A method for specifying time increment control of aircraft flow angle under multiple fault interference

Abstract

The application belongs to the technical field of flight control. The application provides a specified time increment control method for aircraft flow angle under multiple fault interference. The method comprises the following steps: establishing the dynamics equation of the aircraft flow angle and angular rate and converting it into a strict feedback linearization form; performing first-order Taylor expansion on the dynamics equation at the current sampling time to establish a control-oriented incremental model; designing a specified time flow angle controller and a specified time angular rate controller for the aircraft based on the control-oriented incremental model and the specified time theory to achieve specified time control of the aircraft flow angle. The disclosed embodiments have the advantages of simple structure and strong robustness, only require the state differential measurement information and control surface manipulation derivative of the aircraft, effectively reduce the dependence on accurate nonlinear terms in the aircraft model, and can still achieve accurate flow angle control under specified time even under multiple faults.

Description

Technical Field

[0001] The disclosed embodiments relate to the field of flight control technology, and more particularly to a method for controlling an aircraft airflow angle with a specified time increment under multiple fault interferences. Background Art

[0002] An aircraft's airflow angles, namely the angle of attack and angle of sideslip, are crucial parameters for safe flight. Precise and stable control of these angles is crucial for aircraft safety, especially in the event of multiple faults, such as simultaneous jamming and damage of multiple control surfaces, or even asymmetric damage to the wings in extreme environments. These multiple fault interference conditions place higher demands on the designed airflow control law. The control law must not only be robust enough to handle these uncertainties, but also ensure rapid restoration of the aircraft's airflow angle to stability within a short period of time, minimizing the risk of excessive aircraft motion and preventing loss of control.

[0003] Existing approaches address this uncertainty in three main ways. The first involves adaptive control methods, such as model reference adaptation, trajectory linearization, and L1 adaptation. While these methods can mitigate the impact of multiple faults on the aircraft, they are still based on state trim points and are essentially linear time-invariant controllers. When the aircraft state deviates from the controller's designed state point, control performance degrades. Given the large envelope and wide speed range requirements of modern aircraft, this requires the design of multiple adaptive controllers to meet these requirements, increasing the difficulty, time, and labor costs of controller development. The second approach combines nonlinear control methods with adaptive control methods. While adaptive control ensures system robustness, the nonlinear control method is used to eliminate known nonlinear terms in the system, achieving the desired dynamic characteristics. However, in practical applications, the adaptive algorithm's estimation accuracy and sampling time are mismatched. This requires reducing the system sampling time to improve disturbance estimation accuracy, which inevitably increases the computational burden on the system hardware. The third approach utilizes incremental control, which does not explicitly include the complex nonlinear terms of the controlled object and relies solely on sensor measurements and the control effectiveness matrix. This significantly reduces model dependency and offers the advantages of high computational efficiency and minimal computational effort, attracting widespread attention in recent years. When existing incremental control methods are applied to the dynamics of a faulty aircraft, the system often exhibits asymptotic convergence or finite-time convergence. Asymptotic convergence theoretically requires infinite adjustment time, while the convergence time of finite-time control depends on the system's initial state. As the degree of interference from multiple faults on the aircraft increases, the system's convergence time inevitably increases. Fixed-time control methods, on the other hand, suffer from a convergence time that is heavily dependent on the controller's design parameters, limiting their applicability to damaged aircraft.

[0004] Therefore, it is necessary to improve one or more problems in the above related technical solutions.

[0005] It should be noted that this section aims to provide background or context for the technical solutions of the present disclosure stated in the claims. The description herein is not admitted to be prior art merely because it is included in this section. SUMMARY

[0006] The purpose of the embodiments of the present disclosure is to provide a specified time increment control method for aircraft flow angle under multiple fault interference, thereby at least partially overcoming one or more problems caused by the limitations and defects of the related art.

[0007] According to the embodiments of the present disclosure, a specified time increment control method for aircraft flow angle under multiple fault interference is provided, which comprises:

[0008] According to the actual blowing data, a first angular rate differential equation and a first flow angle differential equation of the aircraft are established;

[0009] According to the relationship between the angular rate and the aerodynamic data in the first angular rate differential equation, the aerodynamic data of the aircraft is divided into the control derivative of the aircraft control surface and the aerodynamic parameter related to the state of the aircraft body, to obtain a second angular rate differential equation;

[0010] According to the relationship between the flow angle and the angular rate in the first flow angle differential equation, the aircraft equation is divided into a conversion matrix of the flow angle and the angular rate and a part related to the aerodynamic force of the aircraft, to obtain a second flow angle differential equation;

[0011] The first-order Taylor expansion is performed on the second angular rate differential equation and the second flow angle differential equation, and a control-oriented incremental model is established;

[0012] According to the control-oriented incremental model, a specified time flow angle controller and a specified time angular rate controller of the aircraft are designed;

[0013] According to the specified time flow angle controller, the specified time angular rate controller and the specified time filter, the specified time of the aircraft flow angle is controlled.

[0014] Further, the expression of the first angular rate differential equation is:

[0015]

[0016] The expression of the first flow angle differential equation is:

[0017]

[0018] wherein, is the angle of attack, is the differential of is the differential of is the differential of is the differential of is the differential of is the differential of is the mass of the aircraft, is the airspeed, is the lift, T is the thrust, g is the gravitational acceleration, γ is the path inclination angle, μ is the path roll angle, Y is the side force, θ is the pitch angle, is the roll moment, is the pitch moment, is the yaw moment, , , , , , , , , , , and are the first, second and third moments of inertia of the aircraft about the body axes, is the corresponding product of inertia.

[0019] Further, the expression of the second angular rate differential equation is:

[0020]

[0021] wherein, , is the differential of is the differential of is the aerodynamic parameter related to the state of the aircraft body, is the control derivative of the control surface of the aircraft, , and represent a pair of first and second left elevons on the left side, and represent a pair of first and second right elevons on the right side, represents the rudder, represents the canard.

[0022] Further, the expression of the second flow angle differential equation is:​​​​​​

[0023]

[0024] in, , for The differential of For the aircraft aerodynamic related parts, is the conversion matrix of airflow angle and angular velocity.

[0025] Furthermore, the expression of the control-oriented incremental model is:

[0026]

[0027] Among them, 0 represents the previous sampling moment, for The differential of for The differential of , , , , , represents a higher-order term.

[0028] Furthermore, the expression of the specified time angular rate controller is:

[0029]

[0030] The expression of the specified time airflow angle controller is:

[0031]

[0032] Among them, the first tracking error , the second tracking error , airflow angle reference instruction , for Differential, angular velocity reference instruction , for The differential of is the first parameter to be designed, is the second parameter to be designed, is the third parameter to be designed, is the fourth parameter to be designed, , , is the first time constant, is the second time constant, is the fifth parameter to be designed, is the sixth parameter to be designed, for The differential of for The differential of .

[0033] Furthermore, the expression of the specified time filter is:

[0034]

[0035] Where: is the seventh parameter to be designed, , is the third time constant, It is the eighth parameter to be designed.

[0036] Furthermore, the step of controlling the designated time of the aircraft airflow angle according to the designated time airflow angle controller, the designated time angular rate controller, and the designated time filter includes:

[0037] The airflow angle reference instruction, the angle of attack , the sideslip angle and the roll angle , input into the specified time airflow angle controller to obtain the virtual control input ;

[0038] The virtual control input Input into the specified time filter to obtain the angular velocity reference instruction;

[0039] The angular velocity reference command, the roll angular rate , pitch rate and yaw rate The specified time angular rate controller is input to obtain the airflow angle instruction.

[0040] The technical solutions provided by the embodiments of the present disclosure may have the following beneficial effects:

[0041] In the embodiments disclosed herein, the method for controlling the aircraft's airflow angle at a specified time increment under multiple fault conditions is described above. First, the aircraft's airflow angle and angular rate dynamic equations are established and converted into a strictly feedback linearized form. Second, a first-order Taylor expansion is performed on the dynamic equations at the current sampling time to establish a control-oriented incremental model. Then, based on the control-oriented incremental model and the specified time theory, the aircraft's specified time airflow angle controller and specified time angular rate controller are designed to achieve specified time control of the aircraft's airflow angle. Furthermore, this method has the advantages of a simple structure and strong robustness. It only requires the aircraft's state differential measurement information and control surface manipulation derivatives, effectively reducing its reliance on precise nonlinear terms in the aircraft model. This allows for precise airflow angle control at a specified time even under multiple fault conditions. BRIEF DESCRIPTION OF THE DRAWINGS

[0042] The accompanying drawings are incorporated into and constitute a part of the specification, illustrate embodiments consistent with the present disclosure, and together with the specification, are used to explain the principles of the present disclosure. Obviously, the drawings described below are only some embodiments of the present disclosure, and those skilled in the art can derive other drawings based on these drawings without inventive effort.

[0043] Figure 1 A diagram showing the steps of a method for controlling an aircraft airflow angle at a specified time increment under multiple fault interference in an exemplary embodiment of the present disclosure;

[0044] Figure 2 A specific flow chart showing a method for controlling an aircraft airflow angle at a specified time increment in an exemplary embodiment of the present disclosure;

[0045] Figure 3 The angle of attack instruction in the exemplary embodiment 1 of the present disclosure is shown as , the sideslip angle command is , the roll angle command is Tracking effect when

[0046] Figure 4 It shows that in exemplary embodiment 1 of the present disclosure, at different preset times, 、 and Response situation;

[0047] Figure 5 It shows that in exemplary embodiment 1 of the present disclosure, at different preset times, 、 and Response situation;

[0048] Figure 6 The angle of attack instruction in the exemplary embodiment 2 of the present disclosure is shown as , the sideslip angle command is , the roll angle command is Tracking effect when

[0049] Figure 7 It shows that in exemplary embodiment 2 of the present disclosure, at different preset times, 、 and Response situation;

[0050] Figure 8 It shows that in exemplary embodiment 2 of the present disclosure, at different preset times, 、 and Response situation;

[0051] Figure 9 The angle of attack instruction in the exemplary embodiment 3 of the present disclosure is shown as , the sideslip angle command is , the roll angle command is Tracking effect

[0052] Figure 10 It shows that in exemplary embodiment 3 of the present disclosure, at different preset times, 、 and Response situation;

[0053] Figure 11 It shows that in exemplary embodiment 3 of the present disclosure, at different preset times, 、 and Response situation;

[0054] Figure 12 It is shown that the angle of attack instruction in the exemplary embodiment 4 of the present disclosure is , the sideslip angle command is , the roll angle command is Tracking effect when

[0055] Figure 13 It shows that in exemplary embodiment 4 of the present disclosure, at different preset times, 、 and Response situation;

[0056] Figure 14 It shows that in exemplary embodiment 4 of the present disclosure, at different preset times, and Response situation;

[0057] Figure 15 The angle of attack instruction in exemplary embodiment 5 of the present disclosure is shown as , the sideslip angle command is , the roll angle command is Tracking effect when

[0058] Figure 16 It shows that in exemplary embodiment 5 of the present disclosure, at different preset times, 、 and Response situation;

[0059] Figure 17 It shows that in exemplary embodiment 5 of the present disclosure, at different preset times, 、 and Response situation. DETAILED DESCRIPTION

[0060] Example embodiments will now be described more fully with reference to the accompanying drawings. However, example embodiments can be embodied in many forms and should not be construed as limited to the examples set forth herein; rather, these embodiments are provided so that this disclosure will be thorough and complete and will fully convey the concepts of the example embodiments to those skilled in the art. The described features, structures, or characteristics may be combined in any suitable manner in one or more embodiments.

[0061] In addition, the accompanying drawings are merely schematic illustrations of embodiments of the present disclosure and are not necessarily drawn to scale. Like reference numerals in the figures represent like or similar parts, and thus repeated descriptions thereof will be omitted. Some of the blocks shown in the accompanying drawings are functional entities and do not necessarily correspond to physically or logically separate entities.

[0062] This example embodiment provides a method for controlling the specified time increment of the aircraft airflow angle under multiple fault interference. Figure 1 As shown in , the method for controlling the specified time increment of the aircraft airflow angle under multiple fault interference may include: steps S101 to S106.

[0063] Step S101: establishing a first angular velocity differential equation and a first airflow angle differential equation of the aircraft according to actual wind data;

[0064] Step S102: Based on the relationship between angular rate and aerodynamic data in the first angular rate differential equation, the aircraft aerodynamic data is split into control derivatives of the aircraft control surfaces and aerodynamic parameters related to the aircraft body state, to obtain a second angular rate differential equation;

[0065] Step S103: Based on the relationship between the airflow angle and the angular velocity in the first airflow angle differential equation, the aircraft equation is split into a conversion matrix of the airflow angle and the angular velocity and a part related to the aircraft aerodynamic force, so as to obtain a second airflow angle differential equation;

[0066] Step S104: performing first-order Taylor expansion on the second angular velocity differential equation and the second airflow angle differential equation to establish a control-oriented incremental model;

[0067] Step S105: designing a specified time airflow angle controller and a specified time angular rate controller of the aircraft according to the control-oriented incremental model;

[0068] Step S106: controlling the designated time of the aircraft airflow angle according to the designated time airflow angle controller, the designated time angular rate controller and the designated time filter.

[0069] The above-mentioned specified-time incremental control method for aircraft airflow angles under multiple fault disturbances first establishes the dynamic equations for the aircraft's airflow angle and angular rate and transforms them into a strictly feedback linearized form. Secondly, a first-order Taylor expansion of the dynamic equations at the current sampling time is performed to establish a control-oriented incremental model. Then, based on this control-oriented incremental model and specified-time theory, the aircraft's specified-time airflow angle controller and specified-time angular rate controller are designed to achieve specified-time control of the aircraft's airflow angle. Furthermore, this method offers the advantages of a simple structure and strong robustness. It only requires the aircraft's state differential measurements and control surface manipulation derivatives, effectively reducing its reliance on precise nonlinear terms in the aircraft model. This method can achieve precise airflow angle control at a specified time even under multiple faults.

[0070] Below, we will refer to Figures 1 to 17 Each step of the method for controlling the specified time increment of the aircraft airflow angle under multiple fault interference in this exemplary embodiment is described in more detail.

[0071] In step S101, the aerodynamic data of the aircraft is obtained through actual wind tunnel experiments, and a high-precision six-degree-of-freedom nonlinear model of the aircraft is established. The differential equations of the airflow angular rate and angular velocity of the aircraft are as follows:

[0072] (1)

[0073] (2)

[0074] Where, represents the angle of attack, is the sideslip angle, is the roll angle, is the roll angular rate, is the pitch angular rate, is the yaw angular rate, is the aircraft mass, is the airspeed, is the lift, T is the thrust, g is the acceleration of gravity, γ is the track inclination angle, μ is the track roll angle, Y is the side force, θ is the pitch angle, is the rolling moment, is the pitching moment, is the yaw moment, , , , , , , , , , 、 、 is the moment of inertia of the aircraft around the fuselage axis, The corresponding product of inertia can be obtained through actual measurement.

[0075] In step S102, the relationship between the aerodynamic moment and the aerodynamic moment coefficient in equation (2) is:

[0076] (3)

[0077] Where: is the dynamic pressure, S is the wing area, B is the wing span, C is the wing chord length, is the rolling moment coefficient, is the pitching moment coefficient, is the yaw moment coefficient. Further, the three-axis moment coefficient can be written as:

[0078] (4)

[0079] Where, , , are the aerodynamic parameters related to the aircraft body state, 、 、 The control derivatives generated by the aircraft's control surfaces, , 、 Indicates the left pair of elevons, 、 Indicates the left pair of elevons, Indicates the rudder, Denotes the canard. Considering equation (2), combined with equations (3) and (4), it can be further written as:

[0080] (5)

[0081] In the formula , " " is the vector cross product symbol. Further, equation (5) can be written as:

[0082] (6)

[0083] where , is the aircraft state dependent part, is the aircraft control surface control derivative, and are parameters obtained from actual wind data.

[0084] In step S103, considering the relationship between the flow angle and the angular velocity, equation (1) can be written as:

[0085] (7)

[0086] Further, equation (7) can be written as:

[0087] (8)

[0088] where , is the aircraft aerodynamic force dependent part, is the conversion matrix between the flow angle and the angular velocity.

[0089] In step S104, first order Taylor expansion of the flow angle differential equation (8) and the angular velocity differential equation (6) can be obtained as:

[0090] (9)

[0091] where 0 represents the previous sampling time, represents the high order term. Equation (9) can be further written as:

[0092] (10)

[0093] where is the differential of , is the differential of , , , , , .

[0094] According to the time scale separation principle, and under the assumption that the actuator response is fast enough and the sampling frequency is high, , . Therefore, equation (10) can be written as:

[0095] (11)

[0096] Formula (10) is a control-oriented incremental model.

[0097] In step S105 and step S106, the tracking error is defined as , , , , represents the airflow angle reference command, is the angular velocity reference command. The incremental airflow angle and angular rate controller based on the specified time can be designed as follows:

[0098] (12)

[0099] In the formula, , is the parameter to be designed, , , , is the time constant, , is the parameter to be designed. Preferably, , , , , , , , .

[0100] Considering that in the airflow angle tracking control system based on the incremental backstepping method, the design of the angular rate loop needs the differential signal of the virtual control input , which is easy to cause the problem of differential explosion. Therefore, we design the following specified time filter:

[0101] (13)

[0102] In the formula, , , is the time constant, is the parameter to be designed. Preferably, , , . Then, the airflow angle tracks the airflow angle command within the preset time .

[0103] In summary, the flow chart of the specified time incremental control method of the aircraft airflow angle proposed in the present application is shown in Figure 2 .

[0104] In embodiment 1, as Figures 3 to 5 ​As shown, the simulation results at the specified time of and are shown, the aircraft is trimmed at the height of 1000m and the speed of 60m / s, the trimmed angle of attack is , the trimmed control surface is , , . It can be seen that the angle of attack, sideslip angle and roll angle of the aircraft can track the expected command at the specified time, when the specified time is shorter, the outboard deflection of the aircraft increases, and the actual characteristics of the actuator and the controllable performance of the aircraft are considered in the actual application. Among them,

[0105] Figure 3 are the tracking effects when the angle of attack command is , the sideslip angle command is , and the roll angle command is at the specified time of and .

[0106] Figure 4 are the response conditions of , and under different preset times.

[0107] Figure 5 are the response conditions of , and under different preset times.

[0108] In Example 2, as shown in Figures 6 to 8 , the simulation results at the time of when the control surface is stuck are shown, as shown in Figure 4 (b), 4(c), at 0.5s, and are simultaneously stuck at 0 degrees, causing the roll angle tracking to fluctuate slightly, but the angle of attack and sideslip angle can still accurately track the command in 1s. Among them,

[0109] Figure 6 are the tracking effects when the angle of attack command is , the sideslip angle command is , and the roll angle command is .

[0110] Figure 7 are the response conditions of , and under different preset times.

[0111] Figure 8 are the response conditions of , and .

[0112] In Example 3, as shown in Figures 9 to 11 , the simulation results of the case when the rudder surface is damaged by 50% are shown, compared with the case when the rudder surface is stuck, the rudder effectiveness is reduced by 50%, the rudder output is reduced compared with the normal case, the and are increased to compensate for the loss of moment caused by the reduction of and , the angle of attack, the sideslip angle and the roll angle of the aircraft can achieve accurate tracking of the command, achieving satisfactory results. Among them, are the tracking effects when the angle of attack command is , the sideslip angle command is

[0113] , and the roll angle command is Figure 9 . are the response cases of ,

[0114] and Figure 10 under different preset times. are the response cases of ,

[0115] and Figure 11 under different preset times. are the response cases of ,

[0116] and Figures 12 to 14 under different preset times. In Example 4, as shown in , the simulation results of the case when the right wing surface is damaged by 40% are shown, compared with the case when the rudder surface is stuck and the rudder surface is damaged, the influence of the wing surface damage on the aircraft is significant, especially the response of the roll angle, at the same time, the right wing damage causes to completely fail, the roll angle response has about 1° overshoot, while the angle of attack and the sideslip angle of the aircraft can still achieve accurate tracking of the command, achieving satisfactory results. Among them,

[0117] Figure 12 are the tracking effects when the angle of attack command is , the sideslip angle command is , and the roll angle command is .

[0118] Figure 13 are the response cases of , and under different preset times.

[0119] Figure 14 To set the time at different preset times, and Response situation.

[0120] In Example 5, Figures 15 to 17 As shown, it shows the situation when the aircraft has various failures The simulation results when Stuck at 0 degrees, The rudder efficiency is reduced by 50%, and the right wing is damaged by 40%. Complete failure, the roll angle response has an overshoot of about 4°, and the aircraft's angle of attack and sideslip angle fluctuations increase, but it can still achieve accurate tracking of the command. In short, the algorithm proposed in this application is effective and can achieve the specified time control of the aircraft's airflow angle under multiple failures. Among them,

[0121] Figure 15 The angle of attack command is , the sideslip angle command is , the roll angle command is tracking effect.

[0122] Figure 16 To set the time at different preset times, 、 and Response situation.

[0123] Figure 17 To set the time at different preset times, 、 and Response situation.

[0124] The technical solutions provided by the embodiments of the present disclosure may have the following beneficial effects:

[0125] 1) Strong robustness, which can achieve stable control of aircraft airflow angle under multiple uncertain interferences;

[0126] 2) Fast convergence speed, which can achieve precise control of airflow angle within the time specified by the user;

[0127] 3) Only the airflow angle differential, angular velocity differential signals and control effectiveness matrix are required, without explicitly including complex nonlinear terms, thus reducing the dependence on the precise aircraft model.

[0128] 4) Small amount of calculation, high calculation efficiency, simple structure, and easy application in practical engineering.

[0129] By the above-mentioned specified time increment control method of aircraft flow angle under multiple fault interference, on the one hand, firstly, the aircraft flow angle and angular rate dynamics equation is established and converted into a strict feedback linearization form; secondly, the dynamics equation is first-order Taylor expanded at the current sampling time to establish a control-oriented incremental model; then, based on the control-oriented incremental model and the specified time theory, the specified time flow angle controller and the specified time angular rate controller of the aircraft are designed to realize the specified time control of the aircraft flow angle. On the other hand, the method has the advantages of simple structure and strong robustness, only needs the state differential measurement information and control surface manipulation derivative of the aircraft, effectively reduces the dependence on accurate nonlinear terms in the aircraft model, and accurate flow angle control under specified time can be realized even under multiple faults.

[0130] It should be understood that the terms "center", "longitudinal", "transverse", "length", "width", "thickness", "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", "clockwise", "counterclockwise" and the like in the above description indicate the orientation or positional relationship based on the orientation or positional relationship shown in the drawings, and are only for the convenience of describing the embodiments of the present disclosure and simplifying the description, and do not indicate or imply that the device or element referred to must have a particular orientation, be constructed and operated in a particular orientation, and therefore cannot be understood as a limitation on the embodiments of the present disclosure.

[0131] In addition, the terms "first", "second" are only for descriptive purposes and cannot be understood as indicating or implying relative importance or implicitly indicating the number of the technical features indicated. Therefore, the features defined with "first", "second" can explicitly or implicitly include one or more of the features. In the description of the embodiments of the present disclosure, the meaning of "multiple" is two or more, unless otherwise explicitly specified and limited.

[0132] In the embodiments of the present disclosure, unless otherwise explicitly specified and limited, the terms "mounting", "connection", "connection", "fixing" and the like should be understood in a broad sense, for example, can be fixed connection, can also be detachable connection, or integral; can be mechanical connection, can also be electrical connection; can be directly connected, can also be indirectly connected through an intermediate medium, can be the internal communication of two elements or the interaction relationship between two elements. For those skilled in the art, the specific meaning of the above terms in the present disclosure can be understood according to the specific circumstances.

[0133] In the embodiments of the present disclosure, unless otherwise expressly specified and limited, a first feature being "above" or "below" a second feature may include the first and second features being in direct contact, or may include the first and second features being in contact not directly but through another feature between them. Moreover, a first feature being "above," "above," and "above" a second feature includes the first feature being directly above and obliquely above the second feature, or simply indicates that the first feature is higher in level than the second feature. A first feature being "below," "below," and "below" a second feature includes the first feature being directly below and obliquely below the second feature, or simply indicates that the first feature is lower in level than the second feature.

[0134] In the description of this specification, the reference terms "one embodiment", "some embodiments", "example", "specific example" or "some examples" mean that the specific features, structures, materials or characteristics described in conjunction with the embodiment or example are included in at least one embodiment or example of the present disclosure. In this specification, the schematic expressions of the above terms do not necessarily refer to the same embodiment or example. Moreover, the specific features, structures, materials or characteristics described can be combined in any one or more embodiments or examples in a suitable manner. In addition, those skilled in the art can combine and combine different embodiments or examples described in this specification.

[0135] Those skilled in the art will readily appreciate other embodiments of the present disclosure after considering the specification and practicing the invention disclosed herein. This application is intended to cover any variations, uses, or adaptations of the present disclosure that follow the general principles of the present disclosure and include common knowledge or customary techniques in the art not disclosed herein. The description and examples are to be considered as exemplary only, with the true scope and spirit of the present disclosure being indicated by the appended claims.

Claims

1. A method for controlling aircraft airflow angles at specified time increments under multiple fault disturbances, characterized in that: The method includes: The first angular velocity differential equation and the first airflow angle differential equation of the aircraft are established based on the actual wind data; Based on the relationship between the angular rate and the aerodynamic data in the first angular rate differential equation, the aerodynamic data of the aircraft is split into control derivatives of the aircraft's control surfaces and aerodynamic parameters related to the aircraft's main state, thereby obtaining a second angular rate differential equation; Based on the relationship between the airflow angle and the angular velocity in the first airflow angle differential equation, the aircraft equation is split into a conversion matrix of the airflow angle and the angular velocity and a part related to the aircraft aerodynamic force, so as to obtain a second airflow angle differential equation; Performing a first-order Taylor expansion on the second angular rate differential equation and the second airflow angle differential equation to establish a control-oriented incremental model; Designing a specified time airflow angle controller and a specified time angular rate controller for an aircraft based on the control-oriented incremental model; controlling the designated time of the aircraft airflow angle according to the designated time airflow angle controller, the designated time angular rate controller, and the designated time filter; The expression of the control-oriented incremental model is: Among them, 0 represents the previous sampling moment, for The differential of for The differential of , , , , , represents a higher-order term; The expression of the specified time angular rate controller is: The expression of the specified time airflow angle controller is: Among them, the first tracking error , the second tracking error , airflow angle reference instruction , for Differential, angular velocity reference instruction , for The differential of is the first parameter to be designed, is the second parameter to be designed, is the third parameter to be designed, is the fourth parameter to be designed, , , is the first time constant, is the second time constant, is the fifth parameter to be designed, is the sixth parameter to be designed, for The differential of for The differential of , for The differential of , and Indicates a pair of first left elevon and second left elevon on the left side, and Indicates a pair of first right elevon and second right elevon on the right side, Indicates the rudder, Indicates canard; , for The differential of is the conversion matrix of airflow angle and angular velocity; The expression of the specified time filter is: Where: is the seventh parameter to be designed, , is the third time constant, It is the eighth parameter to be designed.

2. The method for controlling the aircraft airflow angle under multiple fault interferences according to claim 1, characterized in that: The expression of the first angular rate differential equation is: The expression of the first airflow angle differential equation is: in, is the angle of attack, for The differential of is the sideslip angle, The differential of is the roll angle, The differential of is the roll angular rate, The differential of is the pitch angular rate, The differential of is the yaw angular rate, The differential of is the aircraft mass, is the airspeed, is the lift, T is the thrust, g is the acceleration of gravity, γ is the track inclination angle, μ is the track roll angle, Y is the side force, θ is the pitch angle, is the rolling moment, is the pitching moment, is the yaw moment, , , , , , , , , , 、 and are the first moment of inertia, second moment of inertia and third moment of inertia of the aircraft around the fuselage axis, is the corresponding product of inertia.

3. The method for controlling the aircraft airflow angle under multiple fault interferences in a specified time increment according to claim 2, characterized in that: The expression of the second angular rate differential equation is: in, , for The differential of are the aerodynamic parameters related to the aircraft body state, is the control derivative of the aircraft control surface, , and Indicates a pair of first left elevon and second left elevon on the left side, and Indicates a pair of first right elevon and second right elevon on the right side, Indicates the rudder, Indicates canard.

4. The method for controlling the aircraft airflow angle under multiple fault interferences in a specified time increment according to claim 3, characterized in that: The expression of the second airflow angle differential equation is: in, , for The differential of For the aircraft aerodynamic related parts, is the conversion matrix of airflow angle and angular velocity.

5. The method for controlling the aircraft airflow angle under multiple fault interferences in a specified time increment according to claim 4, characterized in that: The step of controlling the designated time of the aircraft airflow angle according to the designated time airflow angle controller, the designated time angular rate controller, and the designated time filter includes: The airflow angle reference instruction, the angle of attack , the sideslip angle and the roll angle , input into the specified time airflow angle controller to obtain the virtual control input ; The virtual control input Input into the specified time filter to obtain the angular velocity reference instruction; The angular velocity reference command, the roll angular rate , pitch angular rate and yaw rate The specified time angular rate controller is input to obtain the airflow angle instruction.

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