Non-singular predetermined time aircraft incremental attitude control method
By designing a non-singular predetermined time incremental attitude control method for aircraft, the problem of rapid and stable attitude control in flight control systems is solved, the robustness and anti-interference ability of the system are realized, singularity is avoided, and stable convergence within a predetermined time is ensured.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- NORTHWESTERN POLYTECHNICAL UNIV
- Filing Date
- 2026-04-14
- Publication Date
- 2026-05-12
AI Technical Summary
Existing flight control systems struggle to achieve fast and stable attitude control when faced with nonlinear dynamics, strong coupling, and external disturbances. Furthermore, the pre-time control method suffers from singularity problems, leading to system instability.
A non-singular predetermined time incremental attitude control method for aircraft is designed. By establishing an aircraft attitude dynamics model, a non-singular predetermined time attitude angle and angular rate controller is designed. Combined with filter output, singular points are avoided, ensuring system stability and fast convergence.
The convergence time of the attitude angle controller can be customized, the system robustness and anti-interference ability are improved, the singularity problem is avoided, the computational burden is reduced, and stable convergence within a predefined time is ensured.
Smart Images
Figure CN122018534A_ABST
Abstract
Description
Technical Field
[0001] This invention relates to the field of attitude control technology, and in particular to a non-singular predetermined time aircraft incremental attitude control method. Background Technology
[0002] Fixed-wing aircraft, with their superior stability and practical performance, play an irreplaceable role in many key fields such as logistics transportation and reconnaissance and search and rescue. Attitude control of fixed-wing aircraft has long faced significant challenges, primarily stemming from the inherent nonlinear dynamics of the aircraft system, strong coupling between multiple channels, and complex control surface allocation issues. Furthermore, it must cope with persistent external airflow disturbances and model uncertainties. Especially in certain special mission scenarios, meeting high maneuverability requirements places even higher demands on the dynamic performance and robustness of the flight control system.
[0003] To address the aforementioned issues, various control strategies have been proposed, such as traditional backstepping control, L1 adaptive control, and incremental dynamic inverse control. However, most of these methods can only guarantee asymptotic convergence of the closed-loop system, and there is still room for improvement in terms of convergence speed and disturbance rejection capability. Therefore, finite-time control methods have been introduced to enhance the robustness of the system and accelerate the convergence process. However, the upper bound of the convergence time of this method heavily depends on the initial state of the system; when the initial value is large, achieving finite-time convergence is very difficult.
[0004] To further overcome the above limitations, fixed-time control strategies have been developed, whose upper bound on convergence time is independent of the system's initial values. However, this upper bound is still affected by system parameters. When model parameters are complex or uncertain, it is difficult to accurately predict the convergence time, which to some extent limits its practical application in flight control systems.
[0005] Preset-time control methods exhibit unique advantages in this context. This method ensures that the tracking error converges to the desired neighborhood within a user-defined timeframe, and this upper bound can be explicitly adjusted independently of the system's initial conditions. More importantly, preset-time control decouples the convergence time from the controller parameters; its settling time depends only on a directly adjustable design parameter, greatly improving its engineering practicality. However, existing preset-time control methods may suffer from singularity problems during the differentiation of virtual control variables, leading to system instability. To avoid singularities, current research often employs piecewise continuous functions or quadratic fractional functions, but this often complicates stability proofs and increases the computational burden on the system.
[0006] Therefore, it is necessary to improve one or more of the problems existing in the above-mentioned related technical solutions.
[0007] It should be noted that this section is intended to provide background or context for the technical solutions of this disclosure as set forth in the claims. The description herein does not constitute an admission that it is prior art simply because it is included in this section. Summary of the Invention
[0008] The purpose of this invention is to provide a non-singular predetermined time aircraft incremental attitude control method, thereby overcoming at least to some extent one or more problems caused by the limitations and defects of related technologies.
[0009] This invention provides a non-singular predetermined time aircraft incremental attitude control method, comprising: S1. Establish an attitude dynamics model of the aircraft based on aerodynamic data, and transform the attitude dynamics model into the angular rate equation of the aircraft. The angular rate equation of the aircraft is a strictly feedback nonlinear form oriented towards control. S2. The angular rate equation of the aircraft is expanded by first-order Taylor to obtain the nonlinear incremental dynamic equation. S3. Based on the attitude angle information and attitude angle command, design the aircraft's predetermined time attitude angle controller, and design a non-singular predefined time convergence form of the virtual attitude angle controller. S4. Based on angular acceleration information, angular velocity information, attitude angle controller and attitude angle virtual controller, design a predetermined time-incremental angular rate controller for the aircraft, and design a non-singular predefined time-converged angular rate virtual controller. S5. Design a non-singular predetermined time filter. Use the derivative of the angular rate command output by the filter and the angular rate command to replace the derivative of the angular rate command of the angular rate controller and the angular rate virtual controller, respectively, to obtain a predetermined time incremental angular rate controller and a predetermined time incremental angular rate virtual controller without singularities.
[0010] In this invention, in S1, the aircraft's attitude dynamics model is as follows:
[0011]
[0012] in, This is the aircraft attitude angle vector. yes The differential, , , These represent the aircraft's pitch angle, roll angle, and yaw angle, respectively. The transformation matrix representing the attitude angle differential and angular rate is given. It is the angular velocity vector. yes The differential, Angular acceleration, , , These represent the roll rate, pitch rate, and yaw rate, respectively. Here is the rotational inertia matrix. This is the aerodynamic moment matrix.
[0013] In this invention, in S1, the angular rate equation of the aircraft is as follows:
[0014] in, It is a nonlinear term. For the control effectiveness matrix, The control performance matrix generated for the control surface. To control the input amount, , , , , , These are the left inner aileron, left outer aileron, right inner aileron, right outer aileron, canard, and rudder. Q For dynamic pressure, For wing area, Let be the diagonal matrix relating the wing's span and mean aerodynamic chord length. and These represent the wing's span and mean aerodynamic chord, respectively. Represents a diagonal matrix. , , These are the rolling moment coefficient, pitching moment coefficient, and yaw moment coefficient generated by the aircraft fuselage, respectively. For the roll damping derivative, For the roll-yaw coupling derivative, For pitch damping derivative, For the yaw-roll coupling derivative, For the heading damping derivative, , , This is airspeed.
[0015] In this invention, the incremental dynamic equation in nonlinear form in S2 is as follows:
[0016] in, The angular acceleration at the sampling time is... Here is the control performance matrix at the sampling time. , For control surface deflection command, The control surface deflection command at the sampling time. For the residual term, The partial derivative of angular acceleration with respect to angular velocity at the sampling time. , The angular velocity at the sampling time, For higher-order minor terms, It gradually decreases as the sampling frequency increases. The modulus satisfies , It is a bounded positive constant.
[0017] In this invention, in S3, the predetermined time attitude angle controller is as follows:
[0018] in, For angular rate commands, This is the attitude angle command. The derivative of the attitude angle command. It is a virtual controller for attitude angles.
[0019] In this invention, the virtual attitude angle controller is as follows:
[0020] in, For attitude angle virtual controller, For attitude angle tracking error, , , For symbolic functions, , , , , , , and These are the control parameters to be designed.
[0021] In this invention, in S4, the predetermined time-incremental angular rate controller is as follows:
[0022] in, For angular rate virtual controller, As shown below:
[0023] in, Indicates angular rate tracking error. , , , , , and These are the positive parameters to be designed.
[0024] In this invention, in S5, the singularity-free predetermined time increment angular rate controller is as follows:
[0025] Predetermined Time Incremental Angular Rate Virtual Controller without Singularity as follows:
[0026] in, This is the angular rate command output by the filter.
[0027] The technical solution provided by this invention may include the following beneficial effects: This invention discloses a non-singular predetermined-time incremental attitude control method for aircraft. By designing the controller as non-singular, the stability of the closed-loop control system is ensured. Furthermore, the incremental nature of the controller effectively improves the system's robustness and anti-interference capability. The convergence time of the attitude angle controller can be user-defined and set by only one explicit adjustable parameter, and the convergence time is unaffected by the system's initial state. The controller has a simple structure, avoiding singularity by eliminating the need for complex piecewise and quadratic fractional functions. The introduction of filters reduces the computational burden of the system and effectively guarantees convergence within the predetermined time. Attached Figure Description
[0028] The accompanying drawings, which are incorporated in and form part of this specification, illustrate embodiments consistent with this disclosure and, together with the description, serve to explain the principles of this disclosure. It is obvious that the drawings described below are merely some embodiments of this disclosure, and those skilled in the art can obtain other drawings based on these drawings without any inventive effort.
[0029] Figure 1 A schematic diagram of a non-singular predetermined time aircraft incremental attitude control method is shown in an exemplary embodiment of this disclosure; Figure 2 Different predefined times are shown in the exemplary embodiments of this disclosure. The aircraft response diagrams under the parameters are shown, where (a) is the roll angle response, (b) is the pitch angle response, and (c) and (d) are the control surface yaw angle responses. Figure 3 The diagram shows the aircraft response at different initial attitude angles provided in the exemplary embodiments of this disclosure, where (a) is the roll angle response, (b) is the pitch angle response, and (c) and (d) are the control surface yaw angle responses. Detailed Implementation
[0030] 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, they are provided so that this disclosure will be more comprehensive and complete, and will fully convey the concept of the exemplary 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.
[0031] Furthermore, the accompanying drawings are merely illustrative diagrams of embodiments of this disclosure and are not necessarily drawn to scale. The same reference numerals in the drawings denote the same or similar parts, and therefore repeated descriptions of them will be omitted. Some block diagrams shown in the drawings are functional entities and do not necessarily correspond to physically or logically independent entities.
[0032] This example implementation provides a non-singular predetermined time aircraft incremental attitude control method, which may include: S1-S5, as follows: S1. Establish an attitude dynamics model of the aircraft based on aerodynamic data, and transform the attitude dynamics model into the angular rate equation of the aircraft. The angular rate equation of the aircraft is a strictly feedback nonlinear form oriented towards control. S2. The angular rate equation of the aircraft is expanded by first-order Taylor to obtain the nonlinear incremental dynamic equation. S3. Based on the attitude angle information and attitude angle command, design the aircraft's predetermined time attitude angle controller, and design a non-singular predefined time convergence form of the virtual attitude angle controller. S4. Based on angular acceleration information, angular velocity information, attitude angle controller and attitude angle virtual controller, design a predetermined time-incremental angular rate controller for the aircraft, and design a non-singular predefined time-converged angular rate virtual controller. S5. Design a non-singular predetermined time filter. Use the derivative of the angular rate command output by the filter and the angular rate command to replace the derivative of the angular rate command of the angular rate controller and the angular rate virtual controller, respectively, to obtain a predetermined time incremental angular rate controller and a predetermined time incremental angular rate virtual controller without singularities.
[0033] In this embodiment, the stability of the closed-loop control system is ensured by designing the controller as non-singular, and the incremental nature of the controller effectively improves the robustness and anti-interference capability of the system. The convergence time of the attitude angle controller can be customized by the user and is set by only one explicit adjustable parameter, and the convergence time is not affected by the initial state of the system. The controller has a simple structure and avoids the singularity of the controller without the need to construct complex piecewise functions and quadratic fractional functions. The introduction of the non-singular predetermined time filter reduces the computational burden of the system and effectively ensures the convergence of the predefined time.
[0034] Please refer to Figure 1 The specific process of each step in the above embodiments will be described below.
[0035] In S1, the aircraft's attitude dynamics model is first established based on aerodynamic data, as shown below: (1) (2) in, This is the aircraft attitude angle vector. yes The differential, , , These represent the aircraft's pitch angle, roll angle, and yaw angle, respectively. The transformation matrix representing the attitude angle differential and angular rate is given. It is the angular velocity vector. yes The differential, Angular acceleration, , , These represent the roll rate, pitch rate, and yaw rate, respectively. Here is the rotational inertia matrix. This is the aerodynamic moment matrix.
[0036] More specifically, ; , , , , These represent the moment of inertia of the roll axis, the moment of inertia of the pitch axis, the moment of inertia of the yaw axis, and the product of the roll-yaw coupling moments of inertia, respectively. , , , These represent the rolling moment, pitching moment, and yaw moment, respectively. Specifically, the details are as follows: (3) in, Q For dynamic pressure, For wing area, Let be the diagonal matrix relating the wing's span and mean aerodynamic chord length. and These represent the wing's span and mean aerodynamic chord, respectively. Represents a diagonal matrix. , , The rolling moment coefficient, pitching moment coefficient, and yaw moment coefficient generated by the aircraft fuselage. For the roll damping derivative, For the roll-yaw coupling derivative, For pitch damping derivative, For the yaw-roll coupling derivative, For the heading damping derivative, The control performance matrix generated for the control surface. , , Airspeed; To control the input amount, , , , , , These are the left inner aileron, left outer aileron, right inner aileron, right outer aileron, canard, and rudder.
[0037] Furthermore, , Each element in the table represents the control performance generated by the corresponding operating surface, obtained through various methods such as wind tunnel testing or CFD aerodynamic simulation. Specifically, For control surfaces The resulting rolling torque controls the derivative. , For control surfaces The resulting pitch moment control derivative, For control surfaces The resulting yaw moment control derivative.
[0038] Then, equation (3) is transformed into the following strictly nonlinear feedback form: (4) in, , .
[0039] In S2, equation (4) is applied at the sampling time. Performing a first-order Taylor expansion, we obtain the following nonlinear incremental dynamic equations: (5) In the formula: The angular acceleration at the sampling time is... Here is the control performance matrix at the sampling time. , For control surface deflection command, The control surface deflection command at the sampling time. For the residual term, The partial derivative of angular acceleration with respect to angular velocity at the sampling time. , The angular velocity at the sampling time, For higher-order minor terms, It gradually decreases as the sampling frequency increases. The modulus satisfies , It is a bounded positive constant.
[0040] In S3, the pre-time attitude angle controller is as follows: (6) in, For angular rate commands, This is the attitude angle command. The derivative of the attitude angle command. It is a virtual controller for attitude angles.
[0041] Furthermore, the virtual controller for the attitude angle is as follows: (7) in, For attitude angle virtual controller, For attitude angle tracking error, , , For symbolic functions, , , , , , , and These are the control parameters to be designed in the attitude angle virtual controller. It should be noted that... , , , None of them have any actual physical meaning.
[0042] In S4, the singularity-free predetermined time incremental angular rate controller is as follows: (8) Among them, the pre-time incremental angular rate virtual controller without singularity As shown below: (9) in, Indicates angular rate tracking error. , , , , , and These are the positive parameters to be designed in the angular rate virtual controller. Similarly, , , , None of them have any actual physical meaning.
[0043] As can be seen from equation (8), the design of the angular rate controller requires the angular rate command to be processed. Take the derivative, for Differentiation yields: (10) In the formula: .
[0044] It is evident that, for Differentiation can lead to the differential explosion problem, increasing the computational burden of the system; and when This can lead to singularities, which may cause the system to become unstable.
[0045] In S5, to avoid direct differentiation of the angular rate command, this application develops a novel predefined time filter to avoid the complexity explosion and singularity problems in S4. The filter design is as follows: (11) In the formula: For the output of the filter, , , , , . , , These are the positive parameters to be designed in the filter.
[0046] Use the filter output and Replacing the angular rate controller (8) and the angular rate virtual controller (9) respectively. and The singularity-free predetermined time angular rate controller is modified as follows: (12) The virtual controller has been modified as follows: (13) Combining controllers (6), (7), (12), (13) and filter (11), the closed-loop attitude control system (1), (5) will operate at a predefined time. Previously stable, .
[0047] Figure 2 Different predefined times provided for embodiments of the present invention The attitude angle response curves are shown, and the parameters of controllers (6), (12) and filter (11) are set as follows: , , , , , , , , Therefore, the predefined time parameters are respectively , , , , The aircraft is trimmed at an altitude of 200m and a speed of 50m / s, with a trim pitch angle of [value missing]. The balancing input is , The pitch angle command is The roll angle command is It can be seen that the aircraft's attitude angles are all within the given predefined time parameters. Previously, the convergence was achieved, but as the convergence time decreased, the attitude angle response exhibited overshoot, and the deflection angle of the control surfaces gradually increased. The system required more control energy to achieve rapid convergence.
[0048] Figure 3 The attitude angle response curves under different initial states provided in this embodiment of the invention are shown below, with the pitch angle command being... The roll angle command is Controller parameters , , , , , , , , Therefore, the predefined time parameters are respectively , Five cases were set up: Case 1 is the initial roll angle. Initial pitch angle Case 2 is the initial roll angle. Initial pitch angle Case 3 is the initial roll angle. Initial pitch angle Case 4 is the initial roll angle. Initial pitch angle Case 5 is the initial roll angle. Initial pitch angle The simulation results show that different initial pitch and roll angles all respond to commands within the user-defined predefined time of 2 seconds, verifying the predefined time performance of this application.
[0049] It should be noted that although several modules of the system for executing actions are mentioned in the detailed description above, this division is not mandatory. In fact, according to embodiments of the present invention, the features and functions of two or more modules described above can be embodied in one module. Conversely, the features and functions of one module described above can be further divided into multiple modules for embodiment. Components shown as modules may or may not be physical units; that is, they may be located in one place or distributed across multiple network units. Some or all of the modules can be selected to achieve the purpose of the present invention according to actual needs. Those skilled in the art can understand and implement this without any inventive effort.
[0050] The embodiments of the present invention have been described above with reference to the accompanying drawings. However, the present invention is not limited to the specific embodiments described above. The specific embodiments described above are merely illustrative and not restrictive. Those skilled in the art can make many other forms under the guidance of the present invention without departing from the spirit and scope of the claims. All of these forms are within the protection scope of the present invention.
Claims
1. A non-singular predetermined time aircraft incremental attitude control method, characterized in that, include: S1. Establish an attitude dynamics model of the aircraft based on aerodynamic data, and transform the attitude dynamics model into the angular rate equation of the aircraft. The angular rate equation of the aircraft is a strictly feedback nonlinear form oriented towards control. S2. The angular rate equation of the aircraft is expanded by first-order Taylor to obtain the nonlinear incremental dynamic equation. S3. Based on the attitude angle information and attitude angle commands, design the aircraft's predetermined time attitude angle controller, and design a non-singular predefined time convergence form of the virtual attitude angle controller. S4. Based on angular acceleration information, angular velocity information, attitude angle controller and attitude angle virtual controller, design a predetermined time-incremental angular rate controller for the aircraft, and design a non-singular predefined time-converged angular rate virtual controller. S5. Design a non-singular predetermined time filter. Use the derivative of the angular rate command output by the filter and the angular rate command to replace the derivative of the angular rate command of the angular rate controller and the angular rate virtual controller, respectively, to obtain a predetermined time incremental angular rate controller and a predetermined time incremental angular rate virtual controller without singularities.
2. The non-singular predetermined time aircraft incremental attitude control method according to claim 1, characterized in that, In S1, the aircraft's attitude dynamics model is as follows: in, This is the aircraft attitude angle vector. yes The differential, , , These represent the aircraft's pitch angle, roll angle, and yaw angle, respectively. The transformation matrix representing the attitude angle differential and angular rate is given. It is the angular velocity vector. yes The differential, Angular acceleration, , , These represent the roll rate, pitch rate, and yaw rate, respectively. Here is the rotational inertia matrix. This is the aerodynamic moment matrix.
3. The non-singular predetermined time aircraft incremental attitude control method according to claim 2, characterized in that, In S1, the equation for the aircraft's angular rate is as follows: in, It is a nonlinear term. For the control effectiveness matrix, The control performance matrix generated for the control surface. To control the input amount, , , , , , These are the left inner aileron, left outer aileron, right inner aileron, right outer aileron, canard, and rudder. Q For dynamic pressure, For wing area, Let be the diagonal matrix relating the wing's span and mean aerodynamic chord length. and These represent the wing's span and mean aerodynamic chord, respectively. Represents a diagonal matrix. , , These are the rolling moment coefficient, pitching moment coefficient, and yaw moment coefficient generated by the aircraft fuselage, respectively. For the roll damping derivative, For the roll-yaw coupling derivative, For pitch damping derivative, For the yaw-roll coupling derivative, For the heading damping derivative, , , This is airspeed.
4. The non-singular predetermined time aircraft incremental attitude control method according to claim 3, characterized in that, In S2, the incremental dynamic equations in nonlinear form are as follows: in, The angular acceleration at the sampling time is... Here is the control performance matrix at the sampling time. , For control surface deflection command, The control surface deflection command at the sampling time. For the residual term, The partial derivative of angular acceleration with respect to angular velocity at the sampling time. , The angular velocity at the sampling time, For higher-order sub-terms, It gradually decreases as the sampling frequency increases. The modulus satisfies , It is a bounded positive constant.
5. The non-singular predetermined time aircraft incremental attitude control method according to claim 4, characterized in that, In S3, the pre-time attitude angle controller is as follows: in, For angular rate commands, This is the attitude angle command. The derivative of the attitude angle command. It is a virtual controller for attitude angles.
6. The non-singular predetermined time aircraft incremental attitude control method according to claim 5, characterized in that, The virtual controller for attitude angles is as follows: in, For attitude angle virtual controller, For attitude angle tracking error, , , For symbolic functions, , , , , , , and These are the control parameters to be designed.
7. The non-singular predetermined time aircraft incremental attitude control method according to claim 6, characterized in that, In S4, the predetermined time incremental angular rate controller is as follows: in, For angular rate virtual controller, As shown below: in, Indicates angular rate tracking error. , , , , , and These are the positive parameters to be designed.
8. The non-singular predetermined time aircraft incremental attitude control method according to claim 7, characterized in that, In S5, the pre-time incremental angular rate controller without singularities is as follows: Predetermined Time Incremental Angular Rate Virtual Controller without Singularity as follows: in, This is the angular rate command output by the filter.