A control method for a high-mobility aircraft

Abstract

The application discloses a control method of a high-maneuverable aircraft, which is designed by applying BLF-Log in the framework of a dynamic surface control method in the process of obtaining a rudder deflection signal, ensures robustness, guarantees that an attack angle command tracking error including an overshoot and a steady-state error for an attack angle command is always kept in a constraint interval, maintains low sensitivity of a control system to noise and isolation, and enables the method to have a faster convergence speed and better convergence quality by using a double-power sliding mode reaching law; the method has strong robustness, can effectively reduce the influence of interference on an aircraft control system, accurately and stably tracks an attack angle command generated by a guidance loop, controls the aircraft to stably and timely generate a required overload, and ensures that the aircraft stably flies and finally implements accurate interception on a target.

Description

Technical Field

[0001] This invention relates to a control method for a precision interception aircraft, specifically a control method for a high-maneuverability aircraft that precisely intercepts high-speed, highly maneuverable targets. Background Technology

[0002] In the field of modern aircraft control, phased-array radar-guided interceptor aircraft are typically capable of precisely intercepting high-speed, highly maneuverable targets. To achieve this goal, interceptor aircraft are required to significantly improve their maneuverability, which places higher demands on aircraft stability control methods. Traditional autopilots are designed based on linear time-invariant models. The selection of operating points ensures that the controller is relatively satisfied with all linearized model operating points. However, the nonlinear terms neglected during linearization will adversely affect the stable operation of the aircraft. When an aircraft performs high-maneuvering flight over a large airspace, it will generate large angles of attack, causing drastic nonlinear changes in the aircraft's aerodynamic parameters, such as air density, atmospheric pressure, temperature, and humidity. At the same time, gusts of wind will also require the nonlinearity of the system to be clearly considered during the controller design process. In addition, when the aircraft adopts a high angle-of-attack flight scheme, the aerodynamic cross-coupling phenomenon between channels will be intensified, making it difficult to establish a relatively accurate mathematical model. Using traditional small-disturbance linearization methods will introduce large errors. In fully strapdown phased array radar-guided aircraft, the guidance loop contains an isolation parasitic loop caused by beam pointing error and radome error, which significantly affects the terminal miss distance of the guidance system. When the autopilot's angle-of-attack response overshoot is large, the control system oscillates more violently, and the impact of the isolation parasitic loop on the guidance and control system will be further aggravated.

[0003] To address the issue of significant nonlinear changes in aerodynamic parameters of phased-array radar-guided interceptor aircraft at high angles of attack, existing technologies present a six-degree-of-freedom dynamic model for flight control at high angles of attack, along with models related to autopilot design. Based on this model, the incremental norm approach is applied to optimize the design of a traditional PI autopilot controller in a nonlinear system, and its performance under nonlinear conditions is analyzed. Although the application of feedback linearization and gain scheduling is one of the most prominent methods for autopilot design, controllers still encounter some unsatisfactory performance when facing highly nonlinear dynamics and large maneuvers. Some researchers have proposed autopilot design based on state-dependent Riccati equations for strongly nonlinear control problems, while others have proposed using nonlinear backstepping control to design robust autopilots. Nevertheless, aerodynamics at high angles of attack is difficult to model accurately, and controllers designed based on accurate mathematical models in autopilot design often fail to deliver satisfactory performance.

[0004] Based on the above problems, the inventors have proposed a control method for highly maneuverable aircraft with dynamic surface nonlinearity based on the double power sliding mode reaching law and the logarithmic barrier Lyapunov function to address the shortcomings of existing research. Summary of the Invention

[0005] To overcome the above problems, the inventors conducted intensive research and designed a control method for high-maneuverability aircraft. In the process of obtaining the rudder deflection signal, this method applies BLF-Log within the framework of dynamic surface control, ensuring robustness while keeping the angle-of-attack command tracking error, including overshoot and steady-state error, within the constraint range. This maintains the control system's low sensitivity to noise and isolation. The double-power sliding mode approach law enables the method to have a faster convergence speed and better convergence quality. This method addresses the trade-off between improving the speed of an aircraft and reducing its sensitivity to noise and isolation. It uses a reduced-order extended state observer to online estimate and compensate for modeling errors such as aerodynamic parameter uncertainties and external disturbances like gusts during flight, enhancing robustness against aerodynamic parameter uncertainties and external disturbances. The method simplifies complexity by converting rudder deflection constraints into control input constraints using an output-input state transition method. This robust method effectively reduces the impact of disturbances on the aircraft control system and accurately and stably tracks the angle-of-attack commands generated by the guidance loop, controlling the aircraft to smoothly and timely generate the required overload, ensuring stable flight and ultimately achieving precise interception of the target, thus completing this invention.

[0006] Specifically, the purpose of this invention is to provide a control method for a high-maneuverability aircraft. In this method, the desired angle of attack output by the guidance system in the aircraft is received in real time, and a rudder deflection signal is obtained and transmitted to the servo motor. The servo motor operates according to the rudder deflection signal, so that the actual angle of attack of the aircraft tracks the desired angle of attack, and the aircraft flies stably until it hits the target.

[0007] The rudder deflection signal is obtained by the following formula (I):

[0008]

[0009] Where u represents the rudder deflection signal;

[0010] G1 represents the design parameters;

[0011] G2 indicates the servo bandwidth;

[0012] This represents an estimate of the total system disturbance.

[0013] This represents the differential value of μ1 obtained through a first-order low-pass filter;

[0014] K2 represents the design parameters;

[0015] S1 represents the tracking error for tracking the desired angle of attack;

[0016] S2 represents the tracking error of the derivative value of μ1 obtained by the first-order low-pass filter;

[0017] Indicates tracking error constraints;

[0018] C stands for array.

[0019] The virtual control quantity μ1 is obtained by differentiating it using a first-order low-pass filter.

[0020] The virtual control quantity μ1 is obtained by the following formula (ii):

[0021]

[0022] in, This represents an estimate of the total system disturbance.

[0023] K1 and K2 represent design parameters, respectively; preferably, K1 > 0 and K2 > 0.

[0024] λ1 and λ2 represent design parameters, preferably λ1 > 1 and 0 < λ2 < 1;

[0025] Indicates the desired angle of attack α c The derivative of .

[0026] The first-order low-pass filter is obtained through the following equation (iii).

[0027]

[0028] Where τ represents the time constant of the first-order low-pass filter.

[0029] S1 is obtained through the following formula (iv), and S2 is obtained through the following formula (v):

[0030] S1=Cx1-α c (Four)

[0031]

[0032] Here, x1 and x2 represent state variables.

[0033] G1 is obtained through equation (vi), and G2 is obtained through equation (vii):

[0034] G2=ω a (seven)

[0035] Where, ω a Indicates the servo bandwidth;

[0036] K α This represents the angular acceleration correction factor for the angle of attack.

[0037] K q This represents the pitch angle and angular acceleration correction factor.

[0038] M represents the Mach number;

[0039] d n d m They represent constants respectively;

[0040] α represents the angle of attack.

[0041] Among them, the estimated value of the total system disturbance Obtained through the following formula (8):

[0042]

[0043] Where ∈2 represents the additional state quantity of the observer;

[0044] ω1 represents the gain of the observer.

[0045] The values ​​of state variable x1 are shown in equation (IX) below, and the values ​​of state variable x2 are shown in equation (X) below:

[0046] x1 = [α q γ M] T (Nine)

[0047] x2=δ (ten)

[0048] Where α represents the angle of attack;

[0049] q represents the pitch rate;

[0050] γ represents the trajectory inclination angle;

[0051] M represents the Mach number;

[0052] δ represents the actual rudder deflection angle.

[0053] The beneficial effects of this invention include:

[0054] (1) The control method for high-maneuverability aircraft provided by the present invention has strong robustness, can effectively reduce the impact of interference on the aircraft control system and can accurately and stably track the angle of attack command generated by the guidance loop, control the aircraft to generate the required overload smoothly and in a timely manner, ensure the stable flight of the aircraft and finally carry out precise interception of the target.

[0055] (2) The control method for high-maneuverability aircraft provided by the present invention significantly improves the control performance of the all-stripper phased array radar-guided aircraft under high angle of attack conditions; taking into full account that the all-stripper phased array radar-guided aircraft has a large stability domain constraining the damping ratio, the logarithmic Lyapunov function is applied to realize the angle of attack tracking error constraint, and the convergence rate is adjusted by the double power sliding mode reaching law to ensure that the aircraft accurately achieves stable tracking of the angle of attack command while satisfying the constraints;

[0056] (3) The control method for high-maneuverability aircraft provided by the present invention uses a reduced-order extended state observer to accurately estimate aerodynamic uncertainties and external disturbances of the system online, which enhances the anti-interference capability of the method. The extended state observer has a simple structure and fewer design parameters, which improves the engineering practicality of the method. Attached Figure Description

[0057] Figure 1 This diagram illustrates the relationship between the system output angle of attack response and the desired angle of attack in Example 1.

[0058] Figure 2 The graph shows the actual rudder deflection angle over time in Example 1;

[0059] Figure 3 This diagram illustrates the relationship between tracking errors when tracking error constraints differ in Example 1.

[0060] Figure 4 The diagram shows the relationship between the output angle of attack response and the desired angle of attack in Comparative Example 1.

[0061] Figure 5 A schematic diagram showing the change of two actual rudder deflection angles over time in Comparative Example 1 is shown;

[0062] Figure 6 A schematic diagram showing the changes of the two actual tracking errors over time in Comparative Example 1 is shown;

[0063] Figure 7 A schematic diagram showing the change of the angle of attack response output in Example 2 over time is shown;

[0064] Figure 8 A schematic diagram showing the actual rudder deflection angle changing over time in Example 2 is shown;

[0065] Figure 9This diagram illustrates how the tracking error changes over time in Example 2. Detailed Implementation

[0066] The present invention will be further described in detail below with reference to the accompanying drawings and embodiments. Through these descriptions, the features and advantages of the present invention will become clearer and more apparent.

[0067] The term “exemplary” as used herein means “serving as an example, embodiment, or illustration.” Any embodiment illustrated herein as “exemplary” is not necessarily to be construed as superior to or better than other embodiments. Although various aspects of embodiments are shown in the accompanying drawings, the drawings are not necessarily drawn to scale unless specifically indicated otherwise.

[0068] According to the present invention, a control method for a high-maneuverability aircraft is provided. In this method, the desired angle of attack output by the guidance system in the aircraft is received in real time, and a rudder deflection signal is obtained and transmitted to the servo motor accordingly. The servo motor operates according to the rudder deflection signal, causing the actual angle of attack of the aircraft to track the desired angle of attack, thus stabilizing the aircraft's flight until it hits the target. The tracking described in this application refers to controlling the actual angle of attack to be as synchronized as possible with the desired angle of attack. The higher the degree of synchronization, the better the tracking effect, the more stable the aircraft's state, and the higher the final hit accuracy. The guidance system in this application receives information transmitted from a phased array radar and obtains the desired angle of attack in real time.

[0069] In a preferred embodiment, the rudder deflection signal is obtained by the following formula (a):

[0070]

[0071] Where u represents the rudder deflection signal;

[0072] G1 represents design parameters and has no specific physical meaning;

[0073] G2 indicates the servo bandwidth;

[0074] This represents an estimate of the total system disturbance.

[0075] This represents the differential value of μ1 obtained through a first-order low-pass filter;

[0076] K2 represents the design parameter; it takes a normal value, and preferably takes the value of 8.

[0077] S1 and S2 both represent state tracking errors; S1 represents the tracking error for tracking the desired angle of attack; S2 represents the tracking error for the differential value of μ1 obtained by the first-order low-pass filter.

[0078] This represents the tracking error constraint; the threshold of this tracking error constraint is set automatically in the simulation system.

[0079] C represents an array; preferably, the value is C = [1 0 0 0].

[0080] Preferably, the virtual control quantity μ1 is differentiated using a first-order low-pass filter to obtain the...

[0081] The virtual control quantity μ1 is obtained by the following formula (ii):

[0082]

[0083] in, This represents an estimate of the total system disturbance.

[0084] K1 represents the design parameter; it takes a normal value, and preferably takes the value of 15.

[0085] λ1 and λ2 represent the design parameters; preferably, the specific values ​​are λ1 = 5 and λ2 = 0.65.

[0086] Indicates the desired angle of attack α c The derivative of .

[0087] In this application, a nonlinear tracking differentiator is used to determine the ideal angle of attack command α. c To obtain by performing tracking differentiation

[0088] The nonlinear tracking differentiator is shown in the following equation:

[0089]

[0090]

[0091] In equation (ii) above of this application, when the state variable is far from the sliding mode, Playing a dominant role, appropriately increasing the values ​​of K1 and λ1 can accelerate the approach speed when the state variables are far from the sliding mode; when the state variables are close to the sliding mode, Playing a dominant role, appropriately increasing the values ​​of K2 and λ2 can accelerate the approach speed when approaching the sliding mode; therefore, this approach law design can ensure that the state variables have high speed and convergence quality in the process of approaching the sliding mode.

[0092] Preferably, the first-order low-pass filter is obtained by the following equation (iii).

[0093] Wherein, τ represents the time constant of the first-order low-pass filter, and its preferred value is 0.001.

[0094] Preferably, the state tracking error S1 is obtained by the following formula (iv), and the state tracking error S2 is obtained by the following formula (v):

[0095] S1=Cx1-α c (Four)

[0096]

[0097] Here, x1 and x2 represent state variables.

[0098] Preferably, G1 is obtained by the following formula (vi), and G2 is obtained by the following formula (vii):

[0099]

[0100] G2=ω a (seven)

[0101] Where, ω a Indicates the servo bandwidth;

[0102] The servo bandwidth in this application is based on the damping ratio ζ and natural frequency ω of the system servo. n The calculation yields the following formula:

[0103]

[0104] K α This represents the angle-of-attack angular acceleration correction coefficient, and its optimal value is K. α =0.7P0S / mV s ;K q This represents the pitch angle angular acceleration correction factor, and its optimal value is K. q =0.7P0SD / I Y P0 represents static pressure, S represents reference area, D represents missile diameter, m represents vehicle mass, and I represents the mass of the projectile. Y V represents the pitching moment. s It indicates the speed of sound.

[0105] M represents the Mach number; it is obtained by measuring a pitot tube and a full-temperature probe.

[0106] d n d m Both represent constants; their optimal value is d. n = -1.948 (rad- 3 ), d m = -11.803 (rad) -1 ).

[0107] α represents the angle of attack, which is calculated from the actual measurement signal of the angular rate gyroscope. In this application, it is desirable to control α to be close to the desired angle of attack. c .

[0108] Preferably, the estimated value of the total system disturbance Obtained through the following formula (8):

[0109]

[0110] Wherein, ∈2 represents the additional state quantity of the observer; the initial value is ∈2(t0)=-ω0x2(t0); in this application, time t0 is the start-up time.

[0111] ω0 and ω1 both represent the observer gain. Preferably, the values ​​of ω0 and ω1 are both 300.

[0112] Preferably, the values ​​of state variable x1 are as shown in equation (ix), and the values ​​of state variable x2 are as shown in equation (x):

[0113] x1 = [α q γ M] T (Nine)

[0114] x2=δ (ten)

[0115] Where α represents the angle of attack; this angle of attack is the actual angle of attack.

[0116] q represents the pitch rate; it is obtained through accelerometer calculation.

[0117] γ represents the trajectory inclination angle; it is obtained through accelerometer calculation.

[0118] M represents the Mach number; it is obtained by measuring a Pitot tube and a full-temperature probe.

[0119] δ represents the actual rudder deflection angle; it is obtained through accelerometer calculation.

[0120] Example 1

[0121] The initial flight Mach number of the aircraft is set to M=3. During the acquisition of state variables x1 and x2, there are sensor measurement errors, resulting in a 30% aerodynamic uncertainty. The rudder deflection signal is obtained in real time using the following formula (I). The servo motor operates according to this rudder deflection signal, enabling the aircraft's actual angle of attack to track the desired angle of attack:

[0122]

[0123] The virtual control quantity μ1 is obtained through the following formula (ii):

[0124]

[0125] The first-order low-pass filter is obtained by the following equation (iii).

[0126]

[0127] S1 is obtained through the following formula (iv), and S2 is obtained through the following formula (v):

[0128] S1=Cx1-α c (Four)

[0129]

[0130] G1 is obtained through the following equation (vi), and G2 is obtained through the following equation (vii):

[0131]

[0132] G2=ω a (seven)

[0133] Estimated value of total system disturbance Obtained through the following formula (8):

[0134]

[0135] The values ​​of state variable x1 are shown in equation (IX), and the values ​​of state variable x2 are shown in equation (X).

[0136] x1 = [α q γ M] T (Nine)

[0137] x2=δ (ten)

[0138] The specific parameter values ​​are shown in the table below:

[0139]

[0140] When the tracking error constraints are respectively and At that time, the simulation results based on the method for obtaining the rudder deflection signal are as follows: Figure 1 , Figure 2 and Figure 3 As shown in the image.

[0141] from Figure 1 As can be seen from the data, when the initial tracking error is large, this method can stably track the angle of attack command output by the guidance system under different tracking constraints.

[0142] Figure 2 The results show that the actual rudder deflection angles all meet the constraints.

[0143] Depend on Figure 3 It can be seen that this method can satisfy stable tracking under different tracking error constraints, with an angle-of-attack response overshoot of 5%, which meets the requirements of the system damping ratio for the stability domain of a fully strapdown phased array radar-guided aircraft.

[0144] Comparative Example 1

[0145] The initial flight Mach number of the aircraft is set to M=3; there are sensor measurement errors in the process of acquiring state variables x1 and x2, resulting in a 30% aerodynamic uncertainty. The rudder deflection signal is obtained in real time through the following formula (XI), and the rudder motor operates according to the rudder deflection signal, so that the actual angle of attack of the aircraft tracks the desired angle of attack: In formula (XI), the rudder deflection signal is obtained by using the double power dynamic surface control method without using the BLF-Log function;

[0146]

[0147] in, This is the output rudder deflection signal;

[0148] There is a 30% aerodynamic uncertainty. To account for gust disturbance, a sinusoidal signal with an amplitude of 3 degrees and a frequency of 0.25 Hz is introduced as an external disturbance into the input channel. The tracking error constraint is set as follows: |δ c | <20°.

[0149] The specific parameter values ​​are the same as in Example 1;

[0150] Simulation results are as follows Figure 4 , Figure 5 , Figure 6 As shown in the image.

[0151] in, Figure 4 , Figure 5 and Figure 6 The DP method in the figure refers to the control result obtained based on Equation (XI) provided in Comparative Example 1, while the BLF-Log method in the figure refers to the control result obtained based on Equation (I) provided in Example 1.

[0152] from Figure 4 It can be seen that both the BLF-Log method and the DP method can accurately and stably estimate the angle of attack command. Figure 5 and Figure 6 It can be seen that the DP method does not constrain the tracking error, and the required rudder deflection angle is smaller than that of the BLF-Log method under the same parameter conditions, resulting in a tracking error that is 67.3% larger than that of the BLF-Log method.

[0153] Example 2

[0154] There is a 30% aerodynamic error and a 20% random disturbance. To account for gust disturbance, a sinusoidal signal with an amplitude of 3deg and a frequency of 0.25Hz is introduced as an external disturbance into the input channel.

[0155] Based on this assumption, the control method in Example 1 was used to perform 100 Monte Carlo simulations, and the results are as follows. Figure 7 , Figure 8 and Figure 9 As shown in the figure, when there are biases and random disturbances in the aerodynamic parameters, the tracking errors of 300 Monte Carlo simulations all meet the tracking error constraint range, and the maximum overshoot is 7.86%. The control method for high-maneuverability aircraft provided in this application has strong robustness.

[0156] The present invention has been described above with reference to preferred embodiments; however, these embodiments are merely exemplary and illustrative. Various substitutions and modifications can be made to the present invention based on these embodiments, all of which fall within the scope of protection of the present invention.

Claims

1. A control method of a high-mobility aircraft, characterized by, In the method, a desired attack angle output by a guidance system in an aircraft is received in real time, and a rudder deflection signal transmitted to a rudder is obtained according to the desired attack angle, the rudder works according to the rudder deflection signal, so that an actual attack angle of the aircraft tracks the desired attack angle, the aircraft flies stably until hitting a target. The rudder deflection signal is obtained by the following formula (I): Wherein, u represents the rudder deflection signal; G1 represents a design parameter, and G2 represents a bandwidth of the rudder; Obtained by the following formula (eight): Wherein, ∈2 represents an additional state quantity of an observer; ω1 represents a gain of the observer; x2 represents a state variable; denotes the derivative of the virtual control quantity μ1 obtained by a first order low pass filter; K2 represents a design parameter; S1 represents a tracking error of the desired attack angle; S2 represents a tracking error of a differential value of μ1 obtained by a first-order low-pass filter; denotes a tracking error constraint; C represents an array; The virtual control quantity μ1 is differentiated by a first-order low-pass filter to obtain the The virtual control quantity μ1 is obtained by the following equation (two): wherein represents an estimate of the total disturbance of the system; K1 and K2 respectively represent design parameters, K1>0 and K2>0; λ1 and λ2 respectively represent design parameters, λ1>1 and 0<λ2<1; denotes the derivative of the desired angle of attack a c of attack a.

2. The control method of the high-maneuverable aircraft according to claim 1, wherein The first order low pass filter obtains the Wherein, τ represents a time constant of the first-order low-pass filter.

3. The control method of the high-maneuverable aircraft according to claim 1, wherein S1 is obtained by the following formula (IV), and S2 is obtained by the following formula (V): S1 = Cx1 - a c (Four) Wherein, x1 and x2 respectively represent state variables.

4. The control method of the high-maneuverable aircraft according to claim 1, wherein G1 is obtained by the following formula (VI), and G2 is obtained by the following formula (VII): G2 = ω a (seven) where ω a represents the bandwidth of the steering engine; K α denotes the angle of attack angular acceleration correction factor; K q represents the pitch angle acceleration correction coefficient; M represents a Mach number; d n , d m represent a constant, respectively; α represents an attack angle.

5. The control method of the high-maneuverable aircraft according to claim 3, wherein The value of the state variable x1 is shown in the following formula (IX), and the value of the state variable x2 is shown in the following formula (X): x1 = [αqγM] T (Nine) x2=δ (X) Wherein, α represents an attack angle; q represents a pitch rate; γ represents a ballistic angle; M represents a Mach number; δ represents an actual rudder deflection angle.

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