An unmanned aerial vehicle terminal vertical overload adaptive control method
By indirectly estimating the angle of attack and overload signals, and combining a first-order differential inertial combination differentiator and inversion method, adaptive control of aircraft overload is realized, solving the problem of balancing speed and stability in aircraft overload control, and improving the stability of control and the convenience of parameter adjustment.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- NAVAL AVIATION UNIV
- Filing Date
- 2023-03-17
- Publication Date
- 2026-07-21
AI Technical Summary
The balance between speed and stability in aircraft overload control is difficult to achieve. Traditional methods are unable to accurately measure angle of attack and angle of attack damping signals, leading to non-minimum phase control problems.
By installing accelerometers and rate gyroscopes to measure vertical acceleration and pitch rate signals, and combining this with aerodynamic parameter estimation, the angle of attack and overload signals are indirectly estimated. A first-order differential inertial combination differentiator and inversion method are designed to achieve adaptive control of angle of attack and overload.
It improves the speed and stability of aircraft overload tracking, solves the problem of low angle of attack measurement accuracy, ensures control stability margin and convenient parameter adjustment, adapts to changes in atmospheric density, and avoids dynamic coefficient drift.
Smart Images
Figure CN116203846B_ABST
Abstract
Description
Technical Field
[0001] This invention relates to the field of aircraft overload control, and more specifically, to an adaptive control method for vertical overload at the end of an unmanned aerial vehicle. Background Technology
[0002] At the terminal stage of an aircraft, significant maneuvering is required, leading many aircraft to switch from attitude control to overload control. Overload control offers superior maneuverability and thus faster response times. However, its stability and stability margin are often difficult to guarantee. Angle-of-attack control, on the other hand, directly affects aircraft stability through its direct parameters of force and torque, making it the fastest and most direct method, while also offering excellent stability margins. However, direct angle-of-attack control is challenging because: firstly, the angle-of-attack signal is difficult to measure accurately; secondly, the angle-of-attack damping signal is even harder to obtain, especially since it is related to the rudder deflection damping signal, which can lead to non-minimum phase control issues. Therefore, traditional methods often employ attitude control to avoid direct angle-of-attack control. Based on the aforementioned background, this invention proposes a method that estimates the angle of attack using angular rate and overload signals instead of directly measuring it, thereby indirectly achieving angle-of-attack tracking stability. Simultaneously, an overload tracking loop is designed with angle-of-attack stability as its core, and the rudder deflection damping signal is reconstructed through inversion. This solves a series of problems faced by traditional control systems and gives the proposed end-of-line overload control excellent stability and stability margin, making this invention highly valuable for engineering applications.
[0003] It should be noted that the information in the background section above is only used to enhance the understanding of the background of the present invention, and therefore may include information that does not constitute prior art known to those skilled in the art. Summary of the Invention
[0004] The purpose of this invention is to provide an adaptive control method for vertical overload at the end of an unmanned aerial vehicle (UAV), thereby overcoming the problem of low dynamic performance and stability margin of overload control of the UAV due to defects in related technologies.
[0005] According to one aspect of the present invention, a method for adaptive control of vertical overload at the end of an unmanned aerial vehicle is provided, comprising the following six steps:
[0006] Step S10: Install an accelerometer on the aircraft to measure the aircraft's vertical acceleration, then divide it by the gravitational acceleration to convert it into a vertical overload signal; then estimate the aircraft's angle of attack signal based on the aircraft's rudder deflection angle signal, vertical overload signal, and estimated aerodynamic parameters of the aircraft to obtain the aircraft's angle of attack estimation signal; then install a rate gyroscope on the aircraft to measure the aircraft's pitch rate signal; then calculate the angle of attack damping estimation signal based on the rudder deflection angle signal, the aircraft's angle of attack estimation signal, and the aircraft's pitch rate signal; then introduce the rudder deflection angle damping signal, superimpose the angle of attack damping estimation signal, and calculate the overload damping estimation signal.
[0007] Step S20: Based on the flight mission of the aircraft, set the desired overload command signal, and then compare it with the vertical overload signal to obtain the overload error signal. Then, integrate the signal to obtain the overload error integral signal. Next, design a first-order differential inertial combination differentiator to obtain the overload error approximate damping signal. Then, superimpose the overload error signal, the overload error integral signal, and the overload damping estimation signal to form the desired angle of attack signal.
[0008] Step S30: The angle-of-attack estimation signal and the angle-of-attack expectation signal of the aircraft are compared to obtain the angle-of-attack error signal; then the angle-of-attack error integral signal is obtained by integration; then the angle-of-attack error damping signal is calculated based on the rudder deflection signal, the pitch rate signal of the aircraft, and the angle-of-attack estimation signal; finally, the angle rate expectation equivalent signal is calculated using the inversion method based on the estimated aerodynamic coefficient of the aircraft, and then the angle-of-attack error signal, the angle-of-attack error integral signal, and the angle-of-attack error damping signal are superimposed to obtain the pitch rate expectation signal.
[0009] Step S40: Compare the expected pitch rate signal with the aircraft pitch rate signal to obtain the pitch rate error signal; then integrate to obtain the pitch rate error integral signal; then calculate the pitch rate error differential estimate signal based on the aircraft angle of attack approximation signal and the aircraft pitch rate signal; then design a first-order differential inertial combination differentiator, input the pitch rate error signal, and obtain the pitch rate error approximation damping signal.
[0010] Step S50: Calculate the aerodynamic adaptive adjustment factor based on the overload error signal, angle of attack error signal, and pitch rate error signal. Then, based on the aerodynamic adaptive adjustment factor, synchronously and adaptively estimate the aircraft's aerodynamic coefficient, aerodynamic control coefficient, and aerodynamic moment coefficient according to their own amplitude.
[0011] Step S60: Based on the aircraft angle of attack estimation signal and the pitch rate error differential estimation signal, calculate the second-order differential estimation signal of the pitch rate error; then, based on the pitch rate error differential estimation signal and the pitch rate error signal, calculate the approximate second-order damping signal of the pitch rate error through a first-order inertial element; then, superimpose the aircraft angle of attack estimation signal and the pitch rate error differential estimation signal to form a rudder deflection damping signal; finally, based on the aircraft angle of attack estimation signal, the aircraft pitch rate signal, the pitch rate error signal, the pitch rate error integral signal, the pitch rate error approximate damping signal, and the pitch rate error differential estimation signal, the final rudder deflection signal is formed and sent to the aircraft servo system to achieve end-stage tracking of the desired overload signal, thereby completing the end-stage pitch channel overload tracking control task of the aircraft.
[0012] In one exemplary embodiment of the present invention, an accelerometer is installed on the aircraft to measure the aircraft's vertical acceleration, which is then divided by gravitational acceleration to convert it into a vertical overload signal. Next, based on the aircraft's rudder deflection angle signal, the vertical overload signal, and estimated aerodynamic parameters, the aircraft's angle of attack signal is estimated to obtain an estimated angle of attack signal. Then, a rate gyroscope is installed on the aircraft to measure the aircraft's pitch rate signal. Finally, based on the rudder deflection angle signal, the estimated angle of attack signal, and the pitch rate signal, an estimated angle of attack damping signal is calculated. Finally, the rudder deflection damping signal is introduced and superimposed with the estimated angle of attack damping signal to calculate the overload damping signal as follows:
[0013]
[0014]
[0015] α gd =ω z -a 34g α g -a 35g δ z ;
[0016]
[0017] Where a y n is the vertical acceleration signal of the aircraft. y The vertical overload signal is represented by g; g is the constant gravitational acceleration; v is the aircraft velocity, which is taken as a constant; δ z The rudder deflection angle signal is initially set to 0, a 35g a 34g This is an estimate of the aerodynamic coefficients of the aircraft, with initial values selected based on wind tunnel test data; α g For the aircraft angle of attack estimation signal, ω zThe pitch rate signal of the aircraft; α gd For the angle-of-attack damping estimation signal; δ zd The rudder deflection damping signal is initially set to 0; n yd This is the overload damping estimation signal.
[0018] In one exemplary embodiment of the present invention, a desired overload command signal is set according to the flight mission of the aircraft, and then compared with the vertical overload signal to obtain an overload error signal. This signal is then integrated to obtain an integrated overload error signal. A first-order differential inertial combination differentiator is then designed to obtain an approximate damping signal for the overload error. Finally, the overload error signal, the integrated overload error signal, and the estimated overload damping signal are superimposed to form a desired angle-of-attack signal, including:
[0019]
[0020] s1=∫e1dt;
[0021]
[0022]
[0023] in e1 is the desired overload command signal; s1 is the overload error signal; dt represents the integration over time; s is the differential operator of the transfer function of the first-order differential inertial combination differentiator; T1 and T2 are the constant time parameters of the first-order differential inertial combination differentiator; e 1d This is an approximate damping signal for overload error; k is the desired angle of attack signal. 11 k 12 k 13 k 14 These are constant control parameters.
[0024] In one exemplary embodiment of the present invention, an angle-of-attack error signal is obtained by comparing the estimated angle-of-attack signal with the desired angle-of-attack signal; then, an integral angle-of-attack error signal is obtained by integration; next, an angle-of-attack error damping signal is calculated based on the rudder deflection signal, the aircraft's pitch rate signal, and the estimated angle-of-attack signal; finally, an equivalent desired pitch rate signal is calculated using an inversion method based on the estimated aerodynamic coefficients of the aircraft; and then, the angle-of-attack error signal, the integral angle-of-attack error signal, and the damped angle-of-attack error signal are superimposed to obtain the desired pitch rate signal.
[0025] e2=α g -α d ;
[0026] s2=∫e2dt;
[0027] e2d =ω z -a 34g α g -a 35g δ z ;
[0028] ω zde =a 34g α g +a 35g δ z ;
[0029] ω zd =ω zde -k 21 e2-k 22 s2-k 23 e 2d ;
[0030] Where e2 is the angle-of-attack error signal; s2 is the integral signal of the angle-of-attack error; e 2d ω is the damping signal for angle of attack error; zde The expected equivalent signal of angular velocity; ω zd The desired pitch rate signal; k 21 k 22 k 23 These are constant control parameters.
[0031] In one exemplary embodiment of the present invention, a pitch rate error signal is obtained by comparing the expected pitch rate signal with the aircraft pitch rate signal; then, integration is performed to obtain the integral pitch rate error signal; next, a differential estimate of the pitch rate error signal is calculated based on the approximate aircraft angle of attack estimate signal and the aircraft pitch rate signal; then, a first-order differential inertial combination differentiator is designed, and the pitch rate error signal is input to obtain the approximate pitch rate error damping signal as follows:
[0032] e3=ω z -ω zd ;
[0033] s3=∫e3dt;
[0034] e 3d =a 24g α g +a 22g ω z ;
[0035]
[0036] Where e3 is the pitch rate error signal; s3 is the pitch rate error integral signal; e 3d The differential estimation signal for pitch rate error; e 3fThis is an approximate damping signal for the pitch rate error; a 24g a 22g This is an estimated value for the aerodynamic moment coefficient of the aircraft, and its initial value is selected based on the wind tunnel test data of the aircraft.
[0037] In one exemplary embodiment of the present invention, the aerodynamic adaptive adjustment factor is calculated based on the overload error signal, angle of attack error signal, and pitch rate error signal. Then, the aerodynamic adaptive adjustment factor is used to synchronously and adaptively estimate the aircraft aerodynamic coefficient, aircraft aerodynamic control coefficient, and aircraft aerodynamic moment coefficient according to their own amplitudes.
[0038]
[0039] a 22g (n+1)=a 22g (n)+e z T;
[0040]
[0041]
[0042]
[0043]
[0044] Where e z a is the aerodynamic adaptive adjustment factor; 25g This is an estimated value for the aerodynamic control coefficient of the aircraft, and its initial value is selected based on wind tunnel test data of the aircraft; k a1 k a2 k a3 k a4 k a5 k a6 ε1, ε2, and ε3 are constant adaptive parameters used to adjust the estimated values of the aircraft's aerodynamic coefficient, aerodynamic control coefficient, and aerodynamic moment coefficient to converge in the same direction; T is a constant integral parameter.
[0045] In one exemplary embodiment of the present invention, the second-order differential estimate signal of the pitch rate error is calculated based on the aircraft angle of attack estimation signal and the pitch rate error differential estimate signal; then, based on the pitch rate error differential estimate signal and the pitch rate error signal, an approximate second-order damping signal of the pitch rate error is calculated through a first-order inertial element; the aircraft angle of attack estimation signal and the pitch rate error differential estimate signal are then superimposed to form a rudder deflection damping signal; finally, the final rudder deflection signal is formed by superimposing the aircraft angle of attack estimation signal, the aircraft pitch rate signal, the pitch rate error signal, the pitch rate error integral signal, the pitch rate error approximate damping signal, and the pitch rate error differential estimate signal as follows:
[0046] e 3dd =a 24g α gd +a 22g e 3d ;
[0047]
[0048] δ zd =-a 24g α gd -(a 22g +k 31 )e 3d -k 32 e3-k 33 e 3dd -k 34 e 3fd ;
[0049] δ z =(-a 24g α g -a 22g ω z )-k 31 e3-k 32 s3-k 33 e 3d -k 34 e 3f ;
[0050] Where e 3dd For the second-order differential estimation signal of pitch rate error, e 3fd The pitch rate error is approximated by a second-order damped signal, δ zd For the rudder deflection damping signal, k 31 k 32 k 33 As a constant control parameter, the rudder deflection angle signal δ is used. zAfter calculating the rudder deflection angle damping signal, and updating the estimated values of the aircraft's aerodynamic coefficient, aerodynamic rudder efficiency coefficient, and aerodynamic torque coefficient, the terminal overload control and tracking tasks of the aircraft can be achieved after the entire closed-loop calculation is completed.
[0051] Beneficial effects
[0052] This invention provides an adaptive control method for vertical overload at the end of an unmanned aerial vehicle (UAV), with the following five main innovations: First, it proposes a method for overload tracking based on angle-of-attack stabilization, which significantly improves the speed and stability of overload tracking. Second, it proposes a method that uses an angle-of-attack estimation signal band instead of angle-of-attack measurement, avoiding the problem of low accuracy in dynamic angle-of-attack measurement. Third, it proposes a method for adaptively adjusting the aerodynamic coefficient based on overload error, angle-of-attack error, and pitch rate error, ensuring that the dynamic coefficient moves synchronously according to its amplitude, better simulating the dynamic coefficient drift problem caused by atmospheric density changes, and avoiding the problem of inconsistent dynamic coefficient drift direction in traditional adaptive methods. Fourth, it proposes a method for estimating and calculating the rudder deflection damping signal, solving the non-minimum phase problem in traditional control, and also achieving accurate estimation of the overload damping signal. The fifth point is that the entire design adopts an inversion method that converts overload into angle of attack, angle of attack driving angular rate, and rudder deflection to achieve angular rate. This has a good progressive physical meaning, makes the selection of control parameters very convenient, and also makes the stability margin of the entire control large.
[0053] It should be understood that the above general description and the following detailed description are exemplary and explanatory only, and are not intended to limit the invention. Attached Figure Description
[0054] The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments consistent with the invention and, together with the description, serve to explain the principles of the invention. It is obvious that the drawings described below are merely some embodiments of the invention, and those skilled in the art can obtain other drawings based on these drawings without any inventive effort.
[0055] Figure 1 This is a flowchart of an adaptive control method for vertical overload at the end of an unmanned aerial vehicle provided by the present invention.
[0056] Figure 2 This is the vertical overload signal curve (unitless) of the method provided in the embodiments of the present invention;
[0057] Figure 3 This is the aircraft angle-of-attack estimation signal curve (unit: degrees) provided by the method in the embodiments of the present invention;
[0058] Figure 4 This is the aircraft pitch rate signal curve (unit: degrees per second) provided by the method in the embodiments of the present invention;
[0059] Figure 5 This is the overload damping estimation signal curve (unitless) of the method provided in the embodiments of the present invention;
[0060] Figure 6 This is the overload error signal curve (unitless) of the method provided in the embodiments of the present invention;
[0061] Figure 7 This is the expected angle of attack signal curve (unit: degrees) of the method provided in the embodiments of the present invention;
[0062] Figure 8 This is the angle-of-attack error signal curve (unit: degrees) of the method provided in the embodiments of the present invention;
[0063] Figure 9 This is the pitch rate error signal curve (unit: degrees) of the method provided in the embodiments of the present invention;
[0064] Figure 10 This is the rudder deflection angle signal curve (unit: degrees) of the method provided in the embodiments of the present invention. Detailed Implementation
[0065] Exemplary embodiments will now be described more fully with reference to the accompanying drawings. However, these exemplary embodiments can be implemented in many forms and should not be construed as limited to the examples set forth herein; rather, these embodiments are provided to make the invention more comprehensive and complete, and to fully convey the concept of the exemplary embodiments to those skilled in the art. The described features, structures, or characteristics can be combined in any suitable manner in one or more embodiments. In the following description, numerous specific details are provided to give a full understanding of embodiments of the invention. However, those skilled in the art will recognize that the technical solutions of the invention may be practiced with one or more of these specific details omitted, or other methods, components, apparatus, steps, etc., may be employed. In other instances, well-known technical solutions are not shown or described in detail to avoid obscuring various aspects of the invention.
[0066] This invention provides a method for adaptive vertical overload control at the end of an aircraft. It measures the aircraft's vertical acceleration information using accelerometers and calculates the vertical overload signal. The angle of attack is estimated by combining this with rudder deflection information. The overload is then compared with the desired overload command to obtain overload error information, which is further integrated. The estimated aerodynamic parameters, rudder deflection damping signal, and pitch rate signal measured by a rate gyroscope are used to estimate the angle of attack damping signal and the overload damping signal. The overload error, integral, and overload damping estimation signals are then combined to obtain the desired angle of attack signal. This desired angle of attack signal is then compared with the estimated angle of attack signal to obtain the angle of attack error information, which is further integrated. The system employs an inversion control method to calculate the desired pitch rate signal, which is then compared with the pitch rate signal to obtain the pitch rate error signal. A first-order differential inertial combination differentiator is then used to calculate the approximate pitch rate error damping signal, and the error signal is used to calculate the aerodynamic adaptive adjustment factor, adaptively adjusting the estimated aerodynamic parameters. The system then superimposes the aircraft angle of attack estimation signal and the pitch rate error differential estimation signal to calculate the rudder deflection damping signal. Finally, the inversion control method is used to combine the pitch rate error signal, the pitch rate error integral signal, and the pitch rate error approximate damping signal to form the final rudder deflection signal, achieving adaptive overload control of the aircraft at the end of the flight path.
[0067] Below, we will combine the appendix Figure 1 The present invention provides a further explanation and description of an adaptive control method for vertical overload at the end of an unmanned aerial vehicle. (Reference) Figure 1 As shown, the adaptive control method for vertical overload at the end of an unmanned aerial vehicle includes the following steps:
[0068] Step S10: Install an accelerometer on the aircraft to measure the aircraft's vertical acceleration, then divide it by the gravitational acceleration to convert it into a vertical overload signal; then estimate the aircraft's angle of attack signal based on the aircraft's rudder deflection angle signal, vertical overload signal, and estimated aerodynamic parameters of the aircraft to obtain the aircraft's angle of attack estimation signal; then install a rate gyroscope on the aircraft to measure the aircraft's pitch rate signal; then calculate the angle of attack damping estimation signal based on the rudder deflection angle signal, the aircraft's angle of attack estimation signal, and the aircraft's pitch rate signal; then introduce the rudder deflection angle damping signal, superimpose the angle of attack damping estimation signal, and calculate the overload damping estimation signal.
[0069] Specifically, this can be broken down into the following four steps. The first step is to install an accelerometer on the aircraft to measure its vertical acceleration, and then divide it by the gravitational acceleration to convert it into a vertical overload signal as follows:
[0070]
[0071] Where a yn is the vertical acceleration signal of the aircraft. y is the vertical overload signal; g is the constant gravitational acceleration.
[0072] The second step involves estimating the aircraft's angle of attack signal based on the aircraft's rudder deflection angle signal, vertical overload signal, and estimated aerodynamic parameters. The resulting estimated angle of attack signal is as follows:
[0073]
[0074] Where v is the velocity of the aircraft, taken as a constant; δ z The rudder deflection angle signal is initially set to 0, a 35g a 34g This is an estimate of the aerodynamic coefficients of the aircraft, with initial values selected based on wind tunnel test data; α g This is used to estimate the angle of attack of an aircraft.
[0075] The third step is to install a rate gyroscope on the aircraft to measure the aircraft's pitch rate signal; then, based on the rudder deflection signal, the estimated angle of attack signal, and the aircraft's pitch rate signal, calculate the estimated angle of attack damping signal as follows:
[0076] α gd =ω z -a 34g α g -a 35g δ z ;
[0077] Where ω z The pitch rate signal of the aircraft; α gd This is the estimated signal for angle-of-attack damping.
[0078] The fourth step involves introducing the rudder deflection angle damping signal, superimposing it with the angle-of-attack damping estimation signal, and calculating the overload damping estimation signal as follows:
[0079]
[0080] Where δ zd The rudder deflection damping signal is initially set to 0; n yd For overload damping estimation signal;
[0081] Step S20: Based on the flight mission of the aircraft, set the desired overload command signal, and then compare it with the vertical overload signal to obtain the overload error signal. Then, integrate the signal to obtain the overload error integral signal. Next, design a first-order differential inertial combination differentiator to obtain the overload error approximate damping signal. Then, superimpose the overload error signal, the overload error integral signal, and the overload damping estimation signal to form the desired angle of attack signal.
[0082] Specifically, this can be broken down into the following three steps. First, based on the aircraft's flight mission, set the desired overload command signal, then compare it with the aforementioned vertical overload signal to obtain the overload error signal, and then integrate it to obtain the integrated overload error signal as follows:
[0083]
[0084] s1=∫e1dt;
[0085] in e1 is the expected overload command signal; s1 is the overload error signal; dt represents the integration over time.
[0086] The second step involves designing a first-order differential inertial combination differentiator to obtain the approximate overload error damping signal as follows:
[0087]
[0088] Where s is the differential operator of the transfer function of the first-order differential inertial combination differentiator, T1 and T2 are the constant time parameters of the first-order differential inertial combination differentiator, and e 1d This is an approximate damping signal for overload error.
[0089] The third step involves superimposing the overload error signal, the integrated overload error signal, and the estimated overload damping signal onto the approximate overload error damping signal to form the desired angle of attack signal, as follows:
[0090]
[0091] in k is the desired angle of attack signal. 11 k 12 k 13 k 14 These are constant control parameters.
[0092] Step S30: The angle-of-attack estimation signal and the angle-of-attack expectation signal of the aircraft are compared to obtain the angle-of-attack error signal; then the angle-of-attack error integral signal is obtained by integration; then the angle-of-attack error damping signal is calculated based on the rudder deflection signal, the pitch rate signal of the aircraft, and the angle-of-attack estimation signal; finally, the angle rate expectation equivalent signal is calculated using the inversion method based on the estimated aerodynamic coefficient of the aircraft, and then the angle-of-attack error signal, the angle-of-attack error integral signal, and the angle-of-attack error damping signal are superimposed to obtain the pitch rate expectation signal.
[0093] Specifically, this can be broken down into the following three steps. First, the estimated angle of attack signal is compared with the expected angle of attack signal to obtain the angle of attack error signal; then, integration is performed to obtain the integral angle of attack error signal as follows:
[0094] e2=α g -α d ;
[0095] s2=∫e2dt;
[0096] Where e2 is the angle of attack error signal; s2 is the angle of attack error integral signal.
[0097] The second step involves calculating the angle-of-attack error damping signal based on the rudder deflection angle signal, the aircraft's pitch rate signal, and the aircraft's angle-of-attack estimation signal, as follows:
[0098] e 2d =ω z -a 34g α g -a 35g δ z ;
[0099] Where e 2d This is the damping signal for angle of attack error.
[0100] The third step involves calculating the expected equivalent signal of the pitch rate based on the estimated aerodynamic coefficients of the aircraft using an inversion method. Then, the angle-of-attack error signal, the integral signal of the angle-of-attack error, and the damped signal of the angle-of-attack error are superimposed to obtain the expected pitch rate signal as follows:
[0101] ω zde =a 34g α g +a 35g δ z ;
[0102] ω zd =ω zde -k 21 e2-k 22 s2-k 23 e 2d ;
[0103] Where ω zde The expected equivalent signal of angular velocity; ω zd The desired pitch rate signal; k 21 k 22 k 23 These are constant control parameters.
[0104] Step S40: Compare the expected pitch rate signal with the aircraft pitch rate signal to obtain the pitch rate error signal; then integrate to obtain the pitch rate error integral signal; then calculate the pitch rate error differential estimate signal based on the aircraft angle of attack approximation signal and the aircraft pitch rate signal; then design a first-order differential inertial combination differentiator, input the pitch rate error signal, and obtain the pitch rate error approximation damping signal.
[0105] Specifically, this can be broken down into the following three steps. First, the expected pitch rate signal is compared with the aircraft's pitch rate signal to obtain the pitch rate error signal; then, integration is performed to obtain the integral pitch rate error signal as follows:
[0106] e3=ω z -ω zd ;
[0107] s3=∫e3dt;
[0108] Where e3 is the pitch rate error signal; s3 is the pitch rate error integral signal.
[0109] The second step involves calculating the differential estimate of the pitch rate error signal based on the approximate estimation signal of the aircraft's angle of attack and the aircraft's pitch rate signal, as follows:
[0110] e 3d =a 24g α g +a 22g ω z ;
[0111] Where e 3d For the differential estimation signal of pitch rate error, a 24g a 22g This is an estimated value for the aerodynamic moment coefficient of the aircraft, and its initial value is selected based on the wind tunnel test data of the aircraft.
[0112] The third step is to design a first-order differential inertial combination differentiator, input the pitch rate error signal, and obtain the approximate damping signal of the pitch rate error as follows:
[0113]
[0114] Where e 3f This is an approximate damping signal for the pitch rate error.
[0115] Step S50: Calculate the aerodynamic adaptive adjustment factor based on the overload error signal, angle of attack error signal, and pitch rate error signal. Then, based on the aerodynamic adaptive adjustment factor, synchronously and adaptively estimate the aircraft's aerodynamic coefficient, aerodynamic control coefficient, and aerodynamic moment coefficient according to their own amplitude.
[0116] Specifically, this can be broken down into the following two steps. The first step is to calculate the aerodynamic adaptive adjustment factor based on the overload error signal, angle-of-attack error signal, and pitch rate error signal, as follows:
[0117]
[0118] Where e z k is the aerodynamic adaptive adjustment factor. a1 k a2 k a3 k a4 k a5 k a6 ε1, ε2, and ε3 are constant adaptive parameters used to adjust the estimated values of the aircraft's aerodynamic coefficient, aerodynamic control coefficient, and aerodynamic torque coefficient to converge in the same direction.
[0119] The second step involves synchronously and adaptively estimating the aircraft's aerodynamic coefficient, aerodynamic control efficiency coefficient, and aerodynamic moment coefficient according to their respective amplitudes, based on the aforementioned aerodynamic adaptive adjustment factor, as follows:
[0120] a 22g (n+1)=a 22g (n)+e z T;
[0121]
[0122]
[0123]
[0124]
[0125] Where a 25g The initial value of the aerodynamic control coefficient of the aircraft is selected based on wind tunnel test data of the aircraft; T is a constant integral parameter; a 24g a 22g This is an estimated value for the aerodynamic moment coefficient of the aircraft, and its initial value is selected based on the wind tunnel test data of the aircraft; a 35g a 34gThis is an estimated value for the aerodynamic coefficient of the aircraft, and its initial value is selected based on the wind tunnel test data of the aircraft.
[0126] Step S60: Based on the aircraft angle of attack estimation signal and the pitch rate error differential estimation signal, calculate the second-order differential estimation signal of the pitch rate error; then, based on the pitch rate error differential estimation signal and the pitch rate error signal, calculate the approximate second-order damping signal of the pitch rate error through a first-order inertial element; then, superimpose the aircraft angle of attack estimation signal and the pitch rate error differential estimation signal to form the rudder deflection damping signal; finally, based on the aircraft angle of attack estimation signal, the aircraft pitch rate signal, the pitch rate error signal, the pitch rate error integral signal, the pitch rate error approximate damping signal, and the pitch rate error differential estimation signal, the final rudder deflection signal is formed and sent to the aircraft servo system to realize the aircraft's tracking of the desired overload signal, thereby completing the task of end-stage overload tracking and control of the aircraft.
[0127] Specifically, it can be broken down into the following five steps. The first step is to calculate the second-order differential estimate of the pitch rate error based on the aircraft angle-of-attack estimation signal and the pitch rate error differential estimation signal, as follows:
[0128] e 3dd =a 24g α gd +a 22g e 3d ;
[0129] Where e 3dd This is the second-order differential estimation signal of the pitch rate error.
[0130] The second step involves using the differential estimation signal of the pitch rate error and the pitch rate error signal to approximate the second-order damping signal of the pitch rate error through a first-order inertial element, as follows:
[0131]
[0132] Where e 3fd The pitch rate error is approximated by a second-order damped signal.
[0133] The third step involves superimposing the approximate second-order damping signal of the pitch rate error onto the aircraft angle-of-attack estimation signal and the differential estimation signal of the pitch rate error to form the rudder deflection damping signal as follows:
[0134] δ zd =-a 24g α gd -(a 22g +k 31 )e 3d -k 32 e3-k 33e 3dd -k 34 e 3fd ;
[0135] Where δ zd This is the rudder deflection damping signal.
[0136] The fourth step involves superimposing the aircraft angle of attack estimation signal, the aircraft pitch rate signal, the pitch rate error signal, the pitch rate error integral signal, the pitch rate error approximate damping signal, and the pitch rate error differential estimation signal to form the final rudder deflection signal, as follows:
[0137] δ z =(-a 24g α g -a 22g ω z )-k 31 e3-k 32 s3-k 33 e 3d -k 34 e 3f ;
[0138] Where k 31 k 32 k 33 k 34 These are constant control parameters.
[0139] The fifth step is to measure the rudder deflection angle signal δ. z The calculation of the rudder deflection angle damping signal completes the estimation of the aircraft's aerodynamic coefficient, aerodynamic rudder efficiency coefficient, and aerodynamic torque coefficient, and updates these values. After the entire closed-loop calculation is completed, the final rudder deflection angle signal is sent to the aircraft's servo system to achieve attitude stability control and tracking.
[0140] Case Implementation and Analysis of Computer Solution Results
[0141] In step S10, g = 9.8, v = 500, and a are selected. 35g The initial value is 0.254, a 34g The initial value is 1.584; the vertical overload signal is obtained by measuring the accelerometer installed on the aircraft, such as... Figure 2 As shown; the aircraft angle of attack estimation signal is obtained as follows: Figure 3 As shown; then, a rate gyroscope is installed on the aircraft to measure the aircraft's pitch rate signal, as shown. Figure 4 As shown; the overload damping estimation signal is as follows: Figure 5 As shown.
[0142] In step S20, select k 11=0.02, k 12 =0.002, k 13 =0.001, the overload error signal is obtained as follows Figure 6 As shown, by selecting T1 = 0.002 and T2 = 0.9, the desired angle of attack signal is obtained as follows: Figure 7 As shown.
[0143] In step S30, select k 21 =4.5, k 22 =0.15, k 23 =0.02, the angle of attack error signal is obtained as follows Figure 8 As shown.
[0144] In step S40, select a 24g The initial value is -193.3, a 22g The initial value is -2.876, and the pitch rate error signal is obtained as follows: Figure 9 As shown. In step S50, select a 25g The initial value is -167.8, T = 0.001.
[0145] In step S60, k is selected. 31 =18, k 32 =0.2, k 33 =0.15, k 34 =0.05, the rudder deflection signal is obtained as follows Figure 10 As shown.
[0146] Depend on Figure 2 It can be seen that the overload instruction can stably approach the desired instruction without overshoot, indicating that its stability margin is large; from Figure 6 It can be seen that the overload tracking error converges rapidly to 0; from Figure 8 It can be seen that the angle-of-attack tracking error also converges quickly to 0; from Figure 10 It can be seen that the maximum value of the rudder deflection signal is less than 10 degrees, and it is larger at the beginning and smaller at the end, indicating good control speed. Furthermore, the overall rudder deflection signal is smooth and oscillating, indicating very stable control, which meets engineering requirements. The overall experimental results show that the control amplitude is reasonable, without sharp oscillations, meeting engineering needs, and the overload response has no overshoot, thus demonstrating that the method has a good stability margin. In conclusion, the method provided by this invention is completely feasible and correct, and has high engineering value.
Claims
1. An adaptive control method for vertical overload at the end of an unmanned aerial vehicle, characterized in that, Includes the following steps: Step S10: Install an accelerometer on the aircraft to measure its vertical acceleration, then divide it by gravitational acceleration to convert it into a vertical overload signal; then estimate the aircraft's angle of attack signal based on the rudder deflection angle signal, the vertical overload signal, and the estimated aerodynamic parameters of the aircraft, obtaining the aircraft's angle of attack estimation signal; then install a rate gyroscope on the aircraft to measure the aircraft's pitch rate signal; then calculate the angle of attack damping estimation signal based on the rudder deflection angle signal, the aircraft's angle of attack estimation signal, and the aircraft's pitch rate signal; finally, introduce the rudder deflection angle damping signal, superimpose the angle of attack damping estimation signal, and calculate the overload damping estimation signal as follows: a gd =ω z -a 34g a g -a 35g d z ; Where a y n is the vertical acceleration signal of the aircraft. y The vertical overload signal is represented by g; g is the constant gravitational acceleration; v is the aircraft velocity, which is taken as a constant; δ z The rudder deflection angle signal is initially set to 0, a 35g a 34g This is an estimate of the aerodynamic coefficients of the aircraft, with initial values selected based on wind tunnel test data; α g For the aircraft angle of attack estimation signal, ω z The pitch rate signal of the aircraft; α gd For the angle-of-attack damping estimation signal; δ zd The rudder deflection damping signal is initially set to 0; n yd For overload damping estimation signal; Step S20: Based on the aircraft's flight mission, a desired overload command signal is set, and then compared with the aforementioned vertical overload signal to obtain an overload error signal. This signal is then integrated to obtain an integrated overload error signal. A first-order differential inertial combination differentiator is then designed to obtain an approximate overload error damping signal. Finally, the overload error signal, the integrated overload error signal, and the estimated overload damping signal are superimposed to form the desired angle of attack signal as follows: s1=∫e1dt; in e1 is the desired overload command signal; s1 is the overload error signal; dt represents the integration over time; s is the differential operator of the transfer function of the first-order differential inertial combination differentiator; T1 and T2 are the constant time parameters of the first-order differential inertial combination differentiator; e 1d This is an approximate damping signal for overload error; k is the desired angle of attack signal. 11 k 12 k 13 k 14 These are constant control parameters; Step S30: Compare the estimated angle-of-attack signal and the desired angle-of-attack signal to obtain the angle-of-attack error signal; then integrate to obtain the integral angle-of-attack error signal; then calculate the angle-of-attack error damping signal based on the rudder deflection signal, the aircraft's pitch rate signal, and the estimated angle-of-attack signal; finally, calculate the equivalent desired angle rate signal using an inversion method based on the estimated aerodynamic coefficients of the aircraft, and then superimpose the angle-of-attack error signal, the integral angle-of-attack error signal, and the damped angle-of-attack error signal to obtain the desired pitch rate signal as follows: e2＝α g -α d ； s2=∫e2dt; e 2d =ω z -a 34g a g -a 35g d z ; oh zde =a 34g a g +a 35g d z ; oh zd =ω zde -k 21 e2-k 22 s2-k 23 e 2d ; Where e2 is the angle-of-attack error signal; s2 is the integral signal of the angle-of-attack error; e 2d ω is the damping signal for angle of attack error; zde The expected equivalent signal of angular velocity; ω zd The desired pitch rate signal; k 21 k 22 k 23 These are constant control parameters; Step S40: Compare the expected pitch rate signal with the aircraft pitch rate signal to obtain the pitch rate error signal; then integrate to obtain the integral pitch rate error signal; next, calculate the differential estimate of the pitch rate error based on the approximate aircraft angle of attack estimate signal and the aircraft pitch rate signal; then design a first-order differential inertial combination differentiator, input the pitch rate error signal, and obtain the approximate pitch rate error damping signal as follows: e3=ω z -oh 2d ; s3=∫e3dt; e 3d =a 24g a g +a 22g oh z ; Where e3 is the pitch rate error signal; s3 is the pitch rate error integral signal; e 3d The differential estimation signal for pitch rate error; e 3f This is an approximate damping signal for the pitch rate error; a 24g a 22g This is an estimated value for the aerodynamic moment coefficient of the aircraft, and its initial value is selected based on the wind tunnel test data of the aircraft. Step S50: Calculate the aerodynamic adaptive adjustment factor based on the overload error signal, angle-of-attack error signal, and pitch rate error signal. Then, based on the aerodynamic adaptive adjustment factor, synchronously and adaptively estimate the aircraft's aerodynamic coefficient, aerodynamic control efficiency coefficient, and aerodynamic moment coefficient according to their own amplitudes as follows: a 22g (n+1)=a 22g (n)+e z T; Where e z a is the aerodynamic adaptive adjustment factor; 25g This is an estimated value for the aerodynamic control coefficient of the aircraft, and its initial value is selected based on wind tunnel test data of the aircraft; k a1 k a2 k a3 k a4 k a5 k a6 ε1, ε2, and ε3 are constant adaptive parameters used to adjust the estimated values of the aircraft's aerodynamic coefficient, aerodynamic control coefficient, and aerodynamic moment coefficient to converge in the same direction; T is a constant integral parameter. Step S60: Based on the aircraft angle of attack estimation signal and the pitch rate error differential estimation signal, calculate the second-order differential estimation signal of the pitch rate error; then, based on the pitch rate error differential estimation signal and the pitch rate error signal, calculate the approximate second-order damping signal of the pitch rate error through a first-order inertial element; then, superimpose the aircraft angle of attack estimation signal and the pitch rate error differential estimation signal to form the rudder deflection damping signal; finally, based on the aircraft angle of attack estimation signal, the aircraft pitch rate signal, the pitch rate error signal, the pitch rate error integral signal, the pitch rate error approximate damping signal, and the pitch rate error differential estimation signal, the final rudder deflection signal is formed and sent to the aircraft servo system to realize the aircraft's tracking of the desired overload signal, thereby completing the aircraft's terminal overload tracking control task as follows: And 3dd =a 24g α gd +a 22g And 3d ; δ zd =-a 24g α gd -(a 22g +k 31 )e 3d -k 32 e3-k 33 e 3dd -k 34 e 3fd ; δ z =(-a 24g α g -the 22g ω z )-s 31 e3s 32 s3-s 33 this 3d -k 34 this 3f ; Where e 3dd For the second-order differential estimation signal of pitch rate error, e 3fd The pitch rate error is approximated by a second-order damped signal, δ zd For the rudder deflection damping signal, k 31 k 32 k 33 k 34 As a constant control parameter, the rudder deflection angle signal δ is used. z After calculating the rudder deflection angle damping signal, and updating the estimated values of the aircraft's aerodynamic coefficient, aerodynamic rudder efficiency coefficient, and aerodynamic torque coefficient, the aircraft can achieve terminal overload adaptive tracking task by completing the entire closed-loop calculation.