Aircraft control system and method using sliding mode control and feedback linearization

Abstract

Aircraft control systems and methods using sliding mode control and feedback linearization. Methods and systems are provided for controlling the bank angle, heading angle, and altitude of an aircraft during flight. The methods and systems disclosed herein utilize sliding mode control and feedback linearization control (nonlinear dynamic control) techniques. The methods and systems can provide autopilot-type functionality that can autonomously execute aggressive maneuvers of the aircraft, as well as more gentle maneuvers.

Description

Technical Field

[0001] The present disclosure relates generally to aircraft and, more particularly, to aircraft control systems and methods. Background Art

[0002] Existing aircraft control systems typically involve tunable parameters whose values ​​must be adjusted based on the aircraft's orientation and flight conditions to achieve desired performance across the aircraft's entire operating envelope. This is often a complex, time-consuming, and expensive task. Some control systems are based on linearization of the aircraft's dynamics around a nominal operating point and therefore do not account for the aircraft's nonlinear dynamics. Summary of the Invention

[0003] In one aspect, the present disclosure describes a method for controlling the bank angle (φ) of an aircraft during flight, the method comprising:

[0004] Receive the commanded bank angle (φ) for the aircraft cmd );

[0005] Calculate the tilt angle error (φ err ), the tilt angle error (φ err ) indicates the aircraft's bank angle (φ) and the command bank angle (φ cmd )

[0006] Use sliding mode control techniques to calculate the target rate of change of the aircraft's bank angle (φ) The input to the sliding mode control technique includes the tilt angle error (φ err );

[0007] The feedback linearization (FL) control technique is used to calculate the target roll rate (P c ), the input to the FL control technique includes the target rate of change of the aircraft's bank angle (φ) as well as

[0008] Use the target body roll rate (P c ) to control one or more actuators of the aircraft.

[0009] FL control techniques may include using the inversion of the relationship between the aircraft's bank angle (φ) and one or more of the aircraft's body angular rates to calculate a target body roll rate (P c ).

[0010] FL control techniques may include using the following formula: To calculate the target body roll rate, Indicates the pitch angle of the aircraft. represents the value of the aircraft's bank angle (φ), A value representing the aircraft's body pitch rate, and A value representing the aircraft's body yaw rate.

[0011] Sliding mode control techniques can include the following: err ), the first threshold (C φ,1 ), the second threshold (C φ,2 ) and the third threshold (C φ,3 ), generating the target rate of change of the tilt angle (φ) When the tilt angle error (φ err ) is greater than the first threshold (C φ,1 ), the target change rate of the tilt angle (φ) is chosen to be substantially equal to the tilt angle saturation rate

[0012] When the tilt angle error (φ err ) is less than the first threshold value (C φ,1 ) and is greater than the second threshold (C φ,2 ), the following formula can be used: To calculate the target rate of change of the tilt angle (φ) Among them, sign(φ err ) is the tilt angle error (φ err ) is a sign function, and k φ Indicates a parameter.

[0013] When the tilt angle error (φ err ) is less than the third threshold value (C φ,3 ), the tilt angle error (φ err ) to select the target rate of change of the tilt angle (φ) using the proportional-integral-derivative control function

[0014] When the tilt angle error (φ err ) is less than the second threshold value (C φ,2 ) and is greater than the third threshold (C φ,3 ), the target rate of change of the tilt angle (φ) can be Select the angle error (φ err ) are proportional.

[0015] When the tilt angle error (φ err ) is less than the second threshold value (C φ,2 ) and is greater than the third threshold (C φ,3 ), the following formula can be used: To calculate the target rate of change of the tilt angle (φ) Among them, k φ Represents a parameter or the described parameter.

[0016] Sliding mode control techniques may include using a sigmoid function as the target rate of change of the tilt angle (φ) and the tilt angle error (φ err ) between them.

[0017] The aircraft may be a blended wing body aircraft.

[0018] Embodiments may include combinations of the above features.

[0019] In another aspect, the present disclosure describes a computer program product for implementing a bank angle control function for an aircraft during flight, the computer program product comprising a non-transitory machine-readable storage medium having program code embodied therein, the program code being readable / executable by a computer, a processor, or a logic circuit to perform the above-described method.

[0020] In another aspect, the present disclosure describes a system for controlling the bank angle (φ) of an aircraft during flight. The system includes:

[0021] One or more computers operatively coupled to receive a commanded bank angle (φ) for the aircraft. cmd ), the one or more computers being configured to:

[0022] Calculate the tilt angle error (φ err ), the tilt angle error (φ err ) indicates the aircraft's bank angle (φ) and the command bank angle (φ cmd )

[0023] Use sliding mode control techniques to calculate the target rate of change of the aircraft's bank angle (φ) The input to the sliding mode control technique includes the tilt angle error (φ err );

[0024] Feedback linearization (FL) control technology is used to calculate the target body roll rate (P c ), the input to the FL control technique includes the target rate of change of the aircraft's bank angle (φ) as well as

[0025] Use the target body roll rate (P c ) to control one or more actuators of the aircraft.

[0026] FL control techniques may include calculating a target body roll rate (P) for the aircraft using the inverse of the relationship between the aircraft's bank angle (φ) and one or more of the aircraft's body angular rates. c ).

[0027] FL control techniques may include using the following formula: To calculate the target body roll rate, Indicates the pitch angle of the aircraft. represents the value of the aircraft's bank angle (φ), A value representing the aircraft's body pitch rate, and A value representing the aircraft's body yaw rate.

[0028] Sliding mode control techniques can include the following: err ), the first threshold (C φ,1 ), the second threshold (C φ,2 ) and the third threshold (C φ,3 ), generating the target rate of change of the tilt angle (φ) When the tilt angle error (φ err ) is greater than the first threshold (C φ,1 ), the target change rate of the tilt angle (φ) is chosen to be substantially equal to the tilt angle saturation rate

[0029] When the tilt angle error (φ err ) is less than the first threshold value (C φ,1 ) and is greater than the second threshold (C φ,2 ), the following formula can be used: To calculate the target rate of change of the tilt angle (φ) Among them, φ err Indicates the tilt angle error, sign(φ err ) is the tilt angle error (φ err ) is a sign function, and k φ Indicates a parameter.

[0030] When the tilt angle error (φ err ) is less than the third threshold value (C φ,3 ), the tilt angle error (φ err ) to select the target rate of change of the tilt angle (φ) using the proportional-integral-derivative control function

[0031] When the tilt angle error (φ err ) is less than the second threshold value (C φ,2 ) and is greater than the third threshold (C φ,3), the target rate of change of the tilt angle (φ) can be It is selected to be substantially equal to the tilt angle error (φ err ) are proportional.

[0032] Tilt angle error (φ err ) is less than the second threshold value (C φ,2 ) and is greater than the third threshold (C φ,3 ), you can use the following formula: To calculate the target rate of change of the tilt angle (φ) Among them, k φ Indicates a parameter.

[0033] Sliding mode control techniques may include using a sigmoid function as the target rate of change of the tilt angle (φ) and the tilt angle error (φ err ) between them.

[0034] Embodiments may include combinations of the above features.

[0035] In another aspect, the present disclosure describes a method for controlling the heading angle (ψ) of an aircraft during flight. The method comprises:

[0036] Receive the command heading angle (ψ cmd );

[0037] Calculate the heading angle error (ψ err ), the heading angle error (ψ err ) indicates the aircraft's heading angle (ψ) and the commanded heading angle (ψ cmd )

[0038] Use sliding mode control techniques to calculate the target rate of change of the aircraft's heading angle (ψ) The input to the sliding mode control technique includes the heading angle error (ψ err );

[0039] Feedback linearization (FL) control technique is used to calculate the aircraft's commanded bank angle (ψ cmd ), the input to the FL control technique includes the target rate of change of the aircraft's heading angle (ψ) as well as

[0040] During flight, the commanded bank angle (φ cmd ) to control one or more actuators of the aircraft.

[0041] FL control techniques may include using the inverse of the relationship between the aircraft's heading angle (ψ) and the aircraft's bank angle (φ) to calculate a commanded bank angle (φ cmd ).

[0042] FL control techniques may include using the following formula: To calculate the command tilt angle (φ cmd ),in, represents the aircraft's true airspeed, and g represents the acceleration due to gravity.

[0043] Sliding mode control techniques can include the following: err ), the first threshold (C ψ,1 ), the second threshold (C ψ,2 ) and the third threshold (C ψ,3 ), generating the target rate of change of the heading angle (ψ) When the heading angle error (ψ err ) is greater than the first threshold (C ψ,1 ), the target change rate of the heading angle (ψ) is chosen to be substantially equal to the heading angle saturation rate

[0044] When the heading angle error (C err ) is less than the first threshold value (C ψ,1 ) and is greater than the second threshold (C ψ,2 ), the following formula can be used: To calculate the target rate of change of heading angle (ψ) Among them, sign(ψ err ) is the heading angle error (ψ err ) is the sign function of .

[0045] When the heading angle error (ψ err ) is less than the third threshold value (C ψ,3 ) can be calculated based on the heading angle error (ψ err ) to select the target rate of change of the heading angle (ψ) using the proportional-integral-derivative control function

[0046] When the heading angle error (ψ err ) is less than the third threshold value (C ψ,3 ), the proportional term in the proportional-integral-derivative control function can be substantially equal to

[0047] When the heading angle error (ψ err ) is less than the second threshold value (C ψ,2 ) and is greater than the third threshold (C ψ,3 ), the target change rate of the heading angle (ψ) can be is chosen to be substantially equal to the heading angle error (ψ err ) are proportional.

[0048] When the heading angle error (ψerr ) is less than the second threshold value (C ψ,2 ) and is greater than the third threshold (C ψ,3 ), you can use the following formula: To calculate the target rate of change of heading angle (ψ)

[0049] Sliding mode control techniques can include using a sigmoid function as the target rate of change of the heading angle (ψ) and heading angle error (ψ err ) between them.

[0050] In some embodiments of the method:

[0051] The FL control technique may be a first FL control technique;

[0052] The sliding mode control technique may be a first sliding mode control technique; and

[0053] During flight, the commanded bank angle (φ cmd ) to control the one or more actuators of the aircraft may include:

[0054] Calculate the tilt angle error (φ cmd ), the tilt angle error (φ cmd ) indicates the aircraft's bank angle (φ) and the command bank angle (φ cmd )

[0055] The second sliding mode control technique is used to calculate the target rate of change of the aircraft's bank angle (φ) The input to the second sliding mode control technique includes the tilt angle error (φ err );

[0056] The second FL control technique is used to calculate the target body roll rate (P c ), the input to the second FL control technique includes the target rate of change of the aircraft's bank angle (φ) as well as

[0057] Use the target body roll rate (P c ) to control the one or more actuators of the aircraft.

[0058] The aircraft may be a blended wing body aircraft.

[0059] Embodiments may include combinations of the above features.

[0060] In another aspect, the present disclosure describes a computer program product for implementing a heading angle control function of an aircraft during flight, the computer program product comprising a non-transitory machine-readable storage medium having program code embodied therein, the program code being readable / executable by a computer, a processor, or a logic circuit to perform the above-described method.

[0061] In another aspect, the present disclosure describes a system for controlling the heading angle (ψ) of an aircraft during flight. The system includes:

[0062] One or more computers operatively coupled to receive a commanded heading angle (ψ cmd ), the one or more computers being configured to:

[0063] Calculate the heading angle error (ψ err ), the heading angle error (ψ err ) indicates the aircraft's heading angle (ψ) and the commanded heading angle (ψ cmd )

[0064] Use sliding mode control techniques to calculate the target rate of change of the aircraft's heading angle (ψ) The input to the sliding mode control technique includes the heading angle error (ψ err );

[0065] Feedback linearization (FL) control technique is used to calculate the aircraft's commanded bank angle (φ cmd ), the input to the FL control technique includes the target rate of change of the aircraft's heading angle (ψ) as well as

[0066] During flight, the commanded bank angle (φ cmd ) to control one or more actuators of the aircraft.

[0067] FL control techniques may include using the inverse of the relationship between the aircraft's heading angle (ψ) and the aircraft's bank angle (φ) to calculate a commanded bank angle (φ cmd ).

[0068] FL control techniques may include using the following formula: To calculate the command tilt angle (φ cmd ),in, represents the aircraft's true airspeed, and g represents the acceleration due to gravity.

[0069] Sliding mode control techniques can include the following: err ), the first threshold (C ψ,1 ), the second threshold (Cψ,2 ) and the third threshold (C ψ,3 ), generating the target rate of change of the heading angle (ψ) When the heading angle error (ψ err ) is greater than the first threshold (C ψ,1 ), the target change rate of the heading angle (ψ) is chosen to be substantially equal to the heading angle saturation rate

[0070] When the heading angle error (ψ err ) is less than the first threshold value (C ψ,1 ) and is greater than the second threshold (C ψ,2 ), the following formula can be used: To calculate the target rate of change of heading angle (ψ) Among them, sign(ψ err ) is the heading angle error (ψ err ) is the sign function of .

[0071] When the heading angle error (ψ err ) is less than the third threshold value (C ψ,3 ) can be calculated based on the heading angle error (ψ err ) to select the target rate of change of the heading angle (ψ) using the proportional-integral-derivative control function

[0072] When the heading angle error (ψ err ) is less than the third threshold value (C ψ,3 ), the proportional term in the proportional-integral-derivative control function can be substantially equal to

[0073] When the heading angle error (ψ err ) is less than the second threshold value (C ψ,2 ) and is greater than the third threshold (C ψ,3 ), the target change rate of the heading angle (ψ) can be is chosen to be substantially equal to the heading angle error (ψ err ) are proportional.

[0074] When the heading angle error (ψ err ) is less than the second threshold value (C ψ,2 ) and is greater than the third threshold (C ψ,3 ), you can use the following formula: To calculate the target rate of change of heading angle (ψ)

[0075] Sliding mode control techniques may include using a sigmoid function as the target rate of change of the heading angle (ψ) and heading angle error (ψ err ) between them.

[0076] In some embodiments of the system:

[0077] The FL control technique may be a first FL control technique;

[0078] The sliding mode control technique may be a first sliding mode control technique; and

[0079] During flight, the commanded bank angle (φ cmd ) to control the one or more actuators of the aircraft may include:

[0080] Calculate the tilt angle error (φ cmd ), the tilt angle error (φ cmd ) indicates the aircraft's bank angle (φ) and the command bank angle (φ cmd )

[0081] The second sliding mode control technique is used to calculate the target rate of change of the aircraft's bank angle (φ) The input to the second sliding mode control technique includes the tilt angle error (φ err );

[0082] The second FL control technique is used to calculate the target body roll rate (P c ), the input to the FL control technique includes the target rate of change of the aircraft's bank angle (φ) as well as

[0083] Use the target body roll rate (P c ) to control the one or more actuators of the aircraft.

[0084] Embodiments may include combinations of the above features.

[0085] In one aspect, the present disclosure describes a method for controlling the altitude (h) of an aircraft during flight. The method comprises:

[0086] Receive the command altitude of the aircraft (h cmd );

[0087] Calculate the height error (h err ), the height error (h err ) indicates the aircraft's altitude (h) and commanded altitude (h cmd )

[0088] Use sliding mode control techniques to calculate the target rate of change of the aircraft's altitude (h) The input to the sliding mode control technique includes the height error (h err );

[0089] A feedback linearization (FL) control technique is used to calculate a value representing a target change in thrust for the aircraft. Inputs to the FL control technique include a target rate of change in the aircraft's altitude (h). as well as

[0090] The values ​​are used during flight to control one or more actuators of the aircraft.

[0091] The value may be a target change in thrust lever angle for the aircraft.

[0092] FL control techniques may include calculating the value using the inverse of the relationship between the aircraft's altitude (h) and the aircraft's airspeed.

[0093] FL control techniques can include:

[0094] Determining an aircraft's true airspeed Rate of change as well as

[0095] Use the following formula: To calculate the target change in thrust (ΔT c ), where g represents the value of gravitational acceleration and m represents the mass of the aircraft.

[0096] Sliding mode control techniques can include err ), the first threshold (C h,1 ), the second threshold (C h,2 ) and the third threshold (C h,3 ), generating the target rate of change of height (h) When the height error (h err ) is greater than the first threshold (C h,1 ), the target change rate of height (h) is chosen to be substantially equal to the high saturation rate

[0097] Height error (h err ) can be less than the first threshold value (C h,1 ) and is greater than the second threshold (C h,2 ), the following formula can be used: To calculate the target rate of change of altitude (h) Among them, sign(h err ) is the height error (h err ) is the sign function of .

[0098] When the height error (h err ) is less than the third threshold value (C h,3 ), we can use the height error (herr ) proportional-integral-derivative control function to select the target rate of change of altitude (h)

[0099] When the height error (h err ) is less than the third threshold value (C h,3 ), the proportional term in the proportional-integral-derivative control function can be substantially equal to

[0100] When the height error (h err ) is less than the second threshold value (C h,2 ) and is greater than the third threshold (C h,3 ), the target change rate of altitude (h) can be Select the height error (h err ) are proportional.

[0101] When the height error (h err ) is less than the second threshold value (C h,2 ) and is greater than the third threshold (C h,3 ), the following formula can be used: To calculate the target rate of change of altitude (h)

[0102] Sliding mode control techniques may include using a sigmoid function as the target rate of change of altitude (h) and height error (h err ) between them.

[0103] The aircraft may be a blended wing body aircraft.

[0104] Embodiments may include combinations of the above features.

[0105] In one aspect, the present disclosure describes a computer program product for implementing an altitude control function for an aircraft during flight, the computer program product comprising a non-transitory machine-readable storage medium having program code embodied therein, the program code being readable / executable by a computer, a processor, or a logic circuit to perform the above-described method.

[0106] In one aspect, the present disclosure describes a system for controlling the altitude (h) of an aircraft during flight. The system includes:

[0107] One or more computers operatively coupled to receive a commanded altitude (h cmd ), the one or more computers being configured to:

[0108] Calculate the height error (h err ), the height error (h err ) indicates the aircraft's altitude (h) and commanded altitude (h cmd )

[0109] Use sliding mode control techniques to calculate the target rate of change of the aircraft's altitude (h) The input to the sliding mode control technique includes the height error (h err );

[0110] Feedback linearization control techniques are used to calculate a value representing the target change in thrust of the aircraft. Inputs to the FL control technique include the target rate of change of the aircraft's altitude (h). as well as

[0111] The values ​​are used during flight to control one or more actuators of the aircraft.

[0112] The value may be a target change in thrust lever angle for the aircraft.

[0113] FL control techniques may include calculating the value using the inverse of the relationship between the aircraft's altitude (h) and the aircraft's airspeed.

[0114] FL control techniques can include:

[0115] Determining an aircraft's true airspeed Rate of change as well as

[0116] Use the following formula: To calculate the target change in thrust rod angle (ΔT c ), where g represents the value of the acceleration due to gravity, and m represents the mass of the aircraft.

[0117] Sliding mode control techniques can include err ), the first threshold (C h,1 ), the second threshold (C h,2 ) and the third threshold (C h,3 ), generating the target rate of change of height (h) So that when the height error (h err ) is greater than the first threshold (C h,1 ), the target change rate of height (h) is chosen to be substantially equal to the high saturation rate

[0118] When the height error (h err ) is less than the first threshold value (C h,1 ) and is greater than the second threshold (C h,2 ), the following formula can be used: To calculate the target rate of change of altitude (h) Among them, sign(h err ) is the height error (h err ) is the sign function of .

[0119] When the height error (h err ) is less than the third threshold value (C h,3 ), based on the height error (h err ) proportional-integral-derivative control function to select the target rate of change of altitude (h)

[0120] When the height error (h err ) is less than the third threshold value (C h,3 ), the proportional term in the proportional-integral-differential control function is substantially equal to

[0121] When the height error (h err ) is less than the second threshold value (C h,2 ) and is greater than the third threshold (C h,3 ), the target change rate of altitude (h) can be Select the height error (h err ) are proportional.

[0122] When the height error (h err ) is less than the second threshold value (C h,2 ) and is greater than the third threshold (C h,3 ), the following formula can be used: To calculate the target rate of change of altitude (h)

[0123] Sliding mode control techniques may include using a sigmoid function as the target rate of change of altitude (h) and height error (h err ) between them.

[0124] Embodiments may include combinations of the above features.

[0125] In one aspect, the present disclosure describes an aircraft comprising a system as described herein. The aircraft may be a blended wing-body aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0126] Reference is now made to the accompanying drawings, in which:

[0127] Figure 1A is a perspective view of an exemplary aircraft including the control system disclosed herein;

[0128] Figure 1B is Figure 1A a graphical representation of an exemplary trajectory executed by an aircraft;

[0129] Figure 1C is Figure 1A a graphical representation of another exemplary trajectory executed by an aircraft;

[0130] Figure 2 yes Figure 1A An exemplary schematic representation of a control system of an aircraft;

[0131] Figure 3 yes Figure 1A Another exemplary schematic representation of a control system of an aircraft;

[0132] Figure 4 yes Figure 2 A schematic representation of an exemplary tilt angle controller of a control system of FIG.

[0133] Figure 5 yes Figure 2 A schematic representation of an exemplary heading angle controller of a control system;

[0134] Figure 6 yes Figure 2 A schematic representation of an exemplary altitude controller of a control system;

[0135] Figure 7 yes Figure 2 Schematic representation of an exemplary internal controller and associated external controller of a control system;

[0136] Figure 8A yes Figure 2 A schematic representation of an exemplary sliding mode controller for a control system of;

[0137] Figure 8B is a diagram of an exemplary sliding mode control mapping between error and target rate;

[0138] Figure 8C is a graph of another exemplary sliding mode control mapping between error and target rate, wherein the mapping utilizes a sigmoid function;

[0139] Figure 9 yes Figure 2 A schematic representation of an exemplary nonlinear dynamic inverse controller of a control system for controlling the bank angle of an aircraft;

[0140] Figure 10 yes Figure 2 A schematic representation of an exemplary nonlinear dynamic inverse controller of a control system for controlling the heading angle of an aircraft;

[0141] Figure 11 yes Figure 2A schematic representation of an exemplary nonlinear dynamic inverse controller of a control system for controlling the altitude of an aircraft;

[0142] Figure 12 yes Figure 2 Another schematic representation of an exemplary tilt angle controller of a control system of;

[0143] Figure 13 yes Figure 2 Another schematic representation of an exemplary heading angle controller of the control system;

[0144] Figure 14 yes Figure 2 Another schematic representation of an exemplary altitude controller of the control system;

[0145] Figure 15 is a flow chart illustrating an exemplary method for controlling the bank angle of an aircraft during flight;

[0146] Figure 16 is a flow chart illustrating an exemplary method for controlling the heading angle of an aircraft during flight; and

[0147] Figure 17 is a flow chart illustrating an exemplary method for controlling the altitude of an aircraft during flight. DETAILED DESCRIPTION

[0148] In various embodiments, the systems and methods described herein can facilitate the development of aircraft control systems for controlling bank angle, heading angle, and / or altitude. In some embodiments, the systems and methods described herein can utilize a combination of sliding mode and / or nonlinear dynamic inversion control techniques.

[0149] Sliding mode control (SMC) is a model-based nonlinear control technique in which the dynamics of an error (i.e., a variable representing the difference between a state variable and its associated command value) can be constrained via a control input that evolves along a sub-manifold (sliding surface) of the corresponding phase space. The sub-manifold can be chosen such that when the error dynamics are constrained to the sub-manifold, the error evolves toward the origin.

[0150] In some embodiments, the methods and systems described herein may include model-based nonlinear control techniques, such as nonlinear dynamic inversion (NDI), in which a first (i.e., inner-loop) control technique, known as feedback linearization (FL), is used to counteract nonlinearities in the dynamics to effectively produce linear dynamics, and a second (i.e., outer-loop) control technique, such as SMC, is used to control the resulting effectively linear dynamics. Using model-based nonlinear control techniques can significantly reduce the need to schedule / adjust one or more control parameters (e.g., controller gains) throughout the aircraft's operating envelope, as the resulting control system can automatically adjust to flight variations by utilizing a dynamic model of the aircraft. Consequently, desired control system performance can be more consistently maintained across the flight envelope (e.g., altitude, speed) and load envelope (e.g., mass, center of gravity (CG) position, inertia). In some embodiments, the methods and systems described herein can facilitate the development of control systems for new aircraft designs by potentially reducing the level of effort required compared to traditional approaches.

[0151] In some embodiments, the methods and systems described herein can also be used to simulate pilot behavior in a drone, allowing such a drone to autonomously follow a trajectory defined by waypoints or perform scripted maneuvers based on time. In some embodiments, the methods and systems described herein can provide an autopilot-type function that can autonomously and relatively smoothly perform both aggressive (e.g., large-angle) maneuvers and more gentle (e.g., small-angle) maneuvers.

[0152] As used herein, the term "substantially" may be applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related.

[0153] Figure 1A is a perspective view of an exemplary aircraft 10 , which may include a system 12 (shown schematically) for controlling some aspects of the operation of the aircraft 10 during flight.

[0154] Figure 1B is a graphical representation of an exemplary two-dimensional projection (ie, a trajectory with respect to two-dimensional geographic coordinates (latitude and longitude)) of a trajectory 11 to be executed by aircraft 10. Trajectory 11 may, for example, be at least partially defined by starting position IC and waypoints WP1-WP4. Figure 1B The trajectory 11 in FIG can be the trajectory of a manned aircraft or an unmanned aircraft.

[0155] Figure 1C is a graphical representation of another exemplary two-dimensional projection of a trajectory 11 to be executed by aircraft 10 . Figure 1CThe trajectory 11 may be defined in part by waypoints or geographic coordinates, which may include latitude, longitude, and altitude. The trajectory 11 may include takeoff, landing, orbit tracking (loitering), test path tracking, waypoint tracking, and / or other types of maneuvers that may be required of the aircraft 10 (blended wing body (BWB) or other types). The systems and methods described herein may facilitate performing such maneuvers and waypoint tracking. The methods and systems described herein may be suitable for facilitating the aircraft 10 to perform two-dimensional or three-dimensional trajectories. Figure 1C The trajectories in can be those of manned or unmanned aircraft.

[0156] refer to Figure 1A The aircraft 10 may be any type of manned or unmanned aircraft (e.g., a drone), such as corporate, private, commercial, and passenger aircraft. For example, the aircraft 10 may be a turboprop aircraft, a (e.g., ultra-long-range) business jet, or a narrow-body twin-engine jetliner. The aircraft 10 may be a fixed-wing aircraft including one or more engines 14. Figure 1A The exemplary aircraft 10 shown in FIG is a blended wing body aircraft. Use of the systems and methods disclosed herein may be particularly advantageous for BWB aircraft having nonlinear flight dynamics that may also change based on changes in the position of the center of gravity CG during flight of such aircraft 10. However, it should be understood that the systems 12 and methods described herein are also applicable to other types of aircraft.

[0157] Aircraft 10 may have a center body 16 having a front end, where a cockpit may be located, and a rear end. Center body 16 may be airfoil-shaped, such as to generate lift. Aircraft 10 may be tailless, but a tail structure may alternatively be provided at the rear end of center body 16. A canard may be provided on the front portion of center body 16. Wings 18 may project laterally from opposite sides of center body 16. Engines 14 may be mounted to the rear end of center body 16. Alternatively or additionally, engines 14 may be mounted to wings 18, or they may be fully or partially embedded within center body 16 or wings 18. BWB aircraft designs are sometimes also referred to as "blended wing-body" aircraft designs. As used herein, the terms "blended wing-body" and "BWB" are intended to encompass designs referred to as "blended wing-body" designs.

[0158] The aircraft 10 may include one or more suitable flight control surfaces 20 configured to interact with air flowing around the aircraft 10 during flight. The control system 12 may be operatively coupled to one or more of these flight control surfaces 20. The one or more flight control surfaces 20 may be removably mounted to the wing 18 and / or other portions of the aircraft 10 and may be configured to cause the aircraft 10 to rotate about an axis B during flight. x 、B y and / or B z Rotation. For example, one or more of the flight control surfaces 20 of the aircraft 10 may be pitch-controlled flight control surfaces (e.g., ailerons, elevons) that are movably mounted to the wings 18 in the case of a BWB aircraft, or to the horizontal stabilizers of the tailplane in the case of a conventional aircraft configuration. Such pitch-controlled flight control surfaces may be considered primary flight control surfaces that allow the aircraft 10 to rotate about a horizontal or lateral axis B during flight. y Move (i.e., rotate). In other words, the movement of the bank angle-controlled flight control surface in flight can cause the aircraft 10 to roll. The flight control surface can be hinged to the trailing edge of the wing 18 or horizontal stabilizer and can move in a controlled manner. The aircraft 10 can also include one or more actuators to adjust the thrust generated by the one or more engines 14. In some embodiments, the engine 14 is a gas turbine engine. In such an embodiment, the actuator can adjust the fuel supply to the engine 14, thereby reducing or increasing the thrust generated by the one or more engines 14. In some embodiments, this thrust adjustment can be used to adjust the orientation of the aircraft 10 during flight.

[0159] The rotation of the aircraft 10 may be measured with respect to different coordinate systems. For example, an earth-fixed coordinate system may have a fixed origin on the earth, an axis G aligned with true north, and an axis G aligned with true north. x , axis G aligned with due east y and the axis G pointing vertically downward z The body-fixed coordinate system may have a fixed origin coincident with the center of gravity CG of the aircraft 10 and may have an axis B oriented toward the nose of the aircraft 10 x , an axis B oriented transversely toward the right wing of the aircraft 10 y , and perpendicular to B x and B y and an axis B oriented toward the belly of the aircraft 10 z The moving earth coordinate system may be defined by translating the earth fixed coordinate system so that its origin coincides with the center of gravity CG of the aircraft 10. The moving earth coordinate system may have an axis H aligned with true north. x , axis H aligned with due east yand the axis H pointing vertically downward z .

[0160] The Earth coordinate system can be moved by rotating it sequentially along three different axes (H x ,H y ,H z ), so that the Earth coordinate system (H x ,H y ,H z ) and the fixed coordinate system of the body (B x ,B y ,B z ) to describe the orientation of the aircraft 10. When this rotation is performed using the yaw-pitch-roll convention described below, three standard aircraft orientation angles can be defined: heading angle ψ, pitch angle θ, and bank angle φ. First, the moving earth coordinate system can be rotated around H by the yaw or heading angle ψ. z Rotate to generate the first intermediate coordinate system (H x(1) ,H y(1) ,H z(1) Then, the first intermediate coordinate system can be rotated around H by the pitch angle θ. y(1) Rotate to generate the second intermediate coordinate system (H x(2) ,H y(2) ,H z(2) Finally, the second intermediate coordinate system can be rotated around H by an inclination angle φ. x(2) Rotate to generate the body fixed coordinate system (B x ,B y ,B z ).

[0161] As the aircraft 10 rotates, the heading angle ψ, the pitch angle θ, and / or the bank angle φ may change, and the aircraft 10 may acquire the heading angle rate Pitch rate and tilt angular rate The corresponding angular rates (with respect to time) indicated by the body roll rate P, body pitch rate Q, and body yaw rate R can be obtained by transforming to a body-fixed coordinate system before evaluating the rate of change with respect to time. The body roll rate P describes the rotation of the aircraft 10 about the axis B. x The body pitch rate Q describes the angular rate of the aircraft 10 about the axis B. y The angular rate of the aircraft 10 about the axis B and the body yaw rate R describe z P, Q and R are collectively referred to as the body angular velocity of the aircraft 10. The heading angular velocity can be defined by the following formula: Pitch rate and tilt angular rate The relationship between the three and the body roll rate P, body pitch rate Q and body yaw rate R:

[0162]

[0163] Figure 2 is an exemplary schematic representation of a control system 12 for aircraft 10. The control system 12 may include one or more pilot or operator input devices 22 for receiving input (e.g., indicative of a desired trajectory for the aircraft 10) from an operator or pilot on or off the aircraft 10. Such input may indicate a desired heading angle (ψ), bank angle (φ), or altitude (h) for the aircraft 10 and may cause the control system 12 to cause movement of one or more flight control surfaces 20 of the aircraft 10 or to cause an increase or decrease in engine thrust, for example, by increasing or decreasing the fuel supply to the burners of the engines 14. The control system 12 may include one or more computers 24 (hereinafter referred to in the singular) operatively coupled to the pilot or operator input devices 22 to receive input signals 26 indicative of the desired commands from the pilot or operator. The pilot or operator input devices 22 may be, for example, a sidestick, a center stick, a control column, a keyboard, or any other combination of input devices configured to receive one or more of a commanded heading angle, bank angle, and altitude. In some embodiments, the systems and methods described herein may also be used with commands generated by an autoflight (e.g., autopilot) system of the aircraft 10, or with trajectory commands (e.g., waypoints) generated based on input from a pilot or other operator remote from the aircraft 10. Thus, the systems and methods disclosed herein may be used for heading, bank, and / or altitude control of either a manned aircraft or an unmanned aircraft (e.g., a drone).

[0164] The computer 24 may include one or more data processors 28 (hereinafter referred to in the singular) and one or more non-transitory memories 30 (i.e., data storage devices) (hereinafter referred to in the singular), the one or more non-transitory memories 30 including machine-readable instructions 32 that are executable by the data processor 28. The instructions 32 can be configured to cause the computer 24 to perform one or more steps in order to implement a computer-implemented process, such that when the instructions 32 are executed by the data processor 28 or other programmable device, the functions / actions specified in the methods described herein are performed. The memory 30 may include any storage means (e.g., device) suitable for retrievably storing machine-readable instructions 32 that are executable by the data processor 28 of the computer 24.

[0165] Various aspects of the present disclosure may be embodied as systems, devices, methods and / or computer program products. Therefore, aspects of the present disclosure may take the form of a complete hardware embodiment, a complete software embodiment, or an embodiment combining software and hardware aspects. In addition, various aspects of the present disclosure may take the form of a computer program product embodied in one or more non-transitory computer-readable media (e.g., memory 30), wherein computer-readable program code is embodied on the non-transitory computer-readable medium. The computer program product may, for example, be executed by a computer 24 so as to perform one or more methods disclosed herein in whole or in part. It should be understood that, based on the present disclosure, a person skilled in the relevant art can easily write a computer program code for implementing the methods disclosed herein.

[0166] Computer 24 may be directly or indirectly operatively coupled to one or more physical actuators 34 (hereinafter referred to in the singular) for controlling the physical actuators 34 associated with one or more flight control surfaces 20 and optionally receiving feedback from the physical actuators 34 and / or coupled to actuators for controlling engine thrust (e.g., fuel pumps, valves). For example, output signals 36 (e.g., command signals) provided by controller 24 may be used to control physical actuators 34. Computer 24 may be considered part of the fly-by-wire flight control system of aircraft 10. For example, computer 24 may be configured to perform functions other than those described herein. In some embodiments, computer 24 may be of the type referred to as the flight control computer (FCC) of aircraft 10. Instructions 32 may be implemented in the form of control laws (CLAWS) in the FCC of aircraft 10. Inputs 26 to computer 24 may also include signals indicating (e.g., sensed or derived) operating parameters (i.e., states) of aircraft 10. Thus, computer 24 may be operatively connected to receive data acquired via one or more sensors 38.

[0167] The various exemplary schematics shown and described herein are not intended to limit the scope of the present disclosure to those described, but rather to provide specific embodiments that may be modified by those skilled in the art within the scope of the present disclosure. The present disclosure encompasses the re-division of subassemblies into potentially different larger or smaller components or subassemblies, or any other reorganization of components and subassemblies.

[0168] Figure 3An exemplary schematic representation of a computer 24 is shown operatively coupled to an actuator 34 and including a plurality of controllers 40-1 through 40-N, where N can be any positive integer. Each controller can receive input signals from a processing block 42 to process signals received from a sensor 38 and / or a state estimation technique 42, where the state estimation technique, such as a Kalman filter or a Luenberger observer or other technique, can involve receiving and / or processing data acquired from the sensor 38. The controllers can be organized such that the output signal of controller 40-(M+1) is used as an input signal to controller 40-M, where M is any integer between 1 and N. In this case, controller 40-M (with respect to controller 40-(M+1)) can be considered an internal controller, while controller 40-(M+1) can be considered an external controller with respect to controller 40-M. Inputs 26 to controller 40 -N may be provided by a pilot / operator, and output signals of controller 40 - 1 may be used as input signals to physical actuators 34 , which may then cause movement of one or more flight control surfaces 20 and / or engine thrust changes.

[0169] Figure 4 A schematic representation of an exemplary control system for controlling the bank angle φ of an aircraft 10 is shown, wherein a bank angle controller 44 defines an outer control loop that receives a commanded bank angle φ from a pilot or operator input device 22 or from another source (e.g., a heading angle controller). cmd The tilt angle controller 44 can generate a target body roll rate P c , then the target body roll rate P c , which is provided as input to a body roll rate controller 46. Both the bank angle controller 44 and the body roll rate controller 46 may receive state information from the sensor signal processing and estimation block 42, which may receive input from one or more sensors 38. The body roll rate controller 46 may then generate a target deflection δ for the control surface 20. The actuator 34 may then deflect the flight control surface 20. The bank angle controller 44 and the body roll rate controller 46 may be part of a larger system that includes one or more computers 24, which may be configured to receive signals from one or more sensors 38 and which may provide input to the actuator 34.

[0170] Figure 5 A schematic diagram of an exemplary control system for controlling a heading angle ψ is shown, wherein a heading angle controller 48 defines an outer control loop that receives a commanded heading angle ψ from a pilot / operator or from another source (e.g., a waypoint tracking controller). cmdThe heading angle controller 48 can generate a commanded tilt angle φ cmd , the command tilt angle φ cmd may be provided as an input to the pitch angle controller 44 which may then generate a target body roll rate P which is provided to the body roll rate controller 46. c The heading angle controller 48, the bank angle controller 44, and the body roll rate controller 46 may all receive state information from the sensor signal processing and estimation block 42, which may receive input from one or more sensors 38. The body roll rate controller 46 may generate a target deflection δ for the control surfaces 20. The actuators 34 may then deflect the one or more flight control surfaces 20. The heading angle controller 48, the bank angle controller 44, and the body roll rate controller 46 may be part of a larger system that includes one or more computers 24 that can be configured to receive signals from the one or more sensors 38 and that may provide input to the actuators 34.

[0171] Figure 6 A schematic diagram of an exemplary control system for controlling the altitude h of an aircraft 10 is shown, wherein an altitude controller 50 defines an outer control loop that receives a commanded altitude h from a pilot / operator or from another source. cmd The altitude controller 50 can generate a target change in thrust ΔT c (e.g., target change in thrust lever angle (TLA)), the target change in thrust ΔT c may be provided as input to a thrust controller 52. Both the altitude controller 50 and the thrust controller 52 may receive state information from a sensor signal processing and estimation block 42, which may receive input from one or more sensors 38. The thrust controller 52 may generate a target change 53 in engine operation to achieve the change indicated by ΔT c Target changes in thrust, such as those indicated by the altitude controller 50 and the thrust controller 52, may be targeted changes in the supply of fuel to the combustors of the gas turbine engines 14 of the aircraft 10. The actuators 34 may implement these targeted changes by regulating (e.g., throttling) the flow of fuel to the engines 14, for example, so that the thrust produced by the engines 14 is changed. The altitude controller 50 and the thrust controller 52 may be part of a larger system that includes one or more computers 24 that can be configured to receive signals from one or more sensors 38 and that can provide input to the actuators 34.

[0172] Figure 7A schematic diagram of an exemplary control architecture including an internal controller 58 and associated external controllers 54 is shown. The external controllers 54 may be a bank angle controller 44, a heading angle controller 48, and / or an altitude controller 50. The internal controllers 58 may be a body roll rate controller 46, a bank angle controller 44, and / or a thrust controller 52, respectively. The external controllers 54 may receive one or more commands 56, such as a commanded bank angle φ, from a pilot / operator or another source. cmd , command heading angle ψ cmd and / or command height h cmd , and receive indications of values ​​of corresponding state variables (such as tilt angle φ, heading angle ψ, altitude h) from the sensor signal processing and / or estimation block 42, which may receive input from one or more sensors 38. In some embodiments, the external controller 54 may receive φ cmd or ψ cmd The output of the sensor signal processing and estimation block 42 may be an input to an upstream controller that provides commands, or it may be an input to an error calculation block 60 within the external controller 54 .

[0173] An error calculation block 60 in the external controller 54 may calculate an error as an indication of the difference between the applicable state variable and the associated command, and then provide the error to a sliding mode control (SMC) block 62. The SMC block 62 may receive the error and, based on the error, generate a target rate of change 55 for the state variable. The rate of change 55 may be an input to a feedback linearization (FL) block 64. The SMC block 62 and the FL block 64 may together form a nonlinear dynamic inverse control (NDIC). The FL block 64 may receive the target rate of change 55 and generate an input 59 (e.g., a target body roll rate P) for the internal controller 58. c , command tilt angle φ cmd and / or target change in thrust ΔT c ). The signals from the internal controller 58 can generate signals that are used to command other controllers operatively coupled to the actuators 34, or to command the actuators 34 directly, to effect a physical change in the state of the aircraft 10 during operation by, for example, deflecting one or more flight control surfaces 20 or increasing the fuel supply to the combustor of a gas turbine engine. The external controller 54 and the internal controller 58 can be part of a larger system that includes one or more computers 24 that can be configured to receive signals from one or more sensors 38 and that can provide input to the actuators 34.

[0174] Figure 8AA schematic diagram of an exemplary SMC block 62 is shown. The SMC block 62 may receive an error from the error calculation block 60. The error may then be processed to obtain a target rate of change 55 of the applicable state variable (e.g., a target rate of change of the bank angle φ of the aircraft 10). Target rate of change of heading angle ψ of aircraft 10 and / or the target rate of change of the altitude h of the aircraft 10 In some embodiments, the error calculation block 60 may be included in the SMC block 62. In some embodiments, other components upstream or downstream of the SMC block 62 may also be included in the SMC block 62. Figure 8A The error handling logic structure is shown within the SMC block 62 in FIG. In some embodiments, the logical conditions of the logic structure can be partially or fully implemented explicitly, for example, by writing software code that implements conditional statements in a computer program. In some embodiments, the logical conditions of the logic structure can be partially or fully implemented implicitly, for example, by using a combination of mechanical and electrical components that implement the logic structure.

[0175] The error received by the SMC block 62 can be passed to a first condition 66 of the logic structure, where if the absolute value of the error is greater than a first threshold 250, the target rate of change can be set equal to the saturation rate. If not, the error can be passed to a second condition 68 of the logic structure. The saturation rate can be user-defined. For example, the saturation rate can be selected to prevent damage to, or extend the useful life of, aircraft components, systems, or structures. In the second condition 68, if the absolute value of the error is less than or equal to the first threshold 250 and greater than a second threshold 252, the target rate of change can be set equal to the product of the offset error and a sign function of the error, where the sign function is -1 when the error is negative, +1 when the error is positive, and 0 when the error is zero. The offset error can be set proportional to the difference between the absolute value of the error and half the second threshold 252. If the second condition 68 is not met, the error may be passed to a third condition 70, where if the absolute value of the error is less than or equal to a second threshold 252 and greater than a third threshold 254, the target rate of change may be set proportional to the error. If the third condition 70 is not met, i.e., the absolute value of the error does not meet any of the previous three conditions and is less than or equal to the third threshold 254, the target rate of change may be set equal to a proportional-integral-derivative (PID) function, i.e., a linear combination of the error, the differential of the error with respect to time, and the integral of the error with respect to time. In some embodiments, the first threshold 250, the second threshold 252, the third threshold 254, and one or more proportionality constants may be selected such that the SMC control block 62 assigns a target rate to each error value in a manner that defines a continuous function that maps the error to a target rate. In some embodiments, the function that maps the error to the target rate may be differentiable. In some embodiments, the function may be differentiable as long as the error is less than the first threshold 250 and / or as long as the error is greater than the first threshold 250.

[0176] Figure 8B is the error sum (such as in Figure 8A , wherein the mapping is differentiable as long as the error is less than a first threshold 250 and as long as the error is greater than the first threshold 250, i.e., the mapping is non-differentiable when the error is equal to the first threshold 250. Between the first threshold 250 and the second threshold 252, the SMC mapping can coincide with a constant braking surface, where the target rate can be equal to the rate produced by the work done by the constant torque or force. Between the second threshold 252 and the third threshold 254, the SMC mapping can coincide with an exponential braking surface, where the target rate is proportional to the error.

[0177] Figure 8Cis a diagram of another exemplary SMC mapping between error and target rate. Figure 8C The SMC mapping utilizes a sigmoid function. In various embodiments, the sigmoid function can be a logistic function, a hyperbolic tangent function (tanh), an inverse tangent function, a Gaussian error function (erf), or any other suitable sigmoid function. The SMC mapping can use a sigmoid function to replace a portion of the above piecewise formula. Since the sigmoid function may be approximately linear for small values ​​of the argument, and as the argument of the function tends to infinity, the sigmoid function asymptotically reaches a maximum value, for example, the standard Gaussian error function asymptotically reaches 1.00, at least a portion of the sigmoid function can be used to specify a target rate based on the error. In some embodiments, for example Figure 8C In the embodiment shown in , a sigmoid function may be used to specify an exponential braking segment between a first threshold 250 and a second threshold 252. Varying the sigmoid function may effectively adjust the slope of the S-shape near the second threshold 252, and thus may allow for greater freedom while maintaining continuity or differentiability of the SMC map across the second threshold 252 in selecting the slope of the exponential braking segment between the second threshold 252 and the third threshold 254.

[0178] In some embodiments, a portion of the exponential braking segment between the second threshold 252 and the third threshold 254, the constant braking segment between the second threshold 252 and the first threshold 250, and the maximum target rate segment greater than the third threshold 250 can all be specified using a sigmoid function. Using the sigmoid function in this manner ensures a smooth transition between the constant braking segment and the maximum target rate segment. The sigmoid function can be modified based on the maximum target rate and the three thresholds 250, 252, and 254 to achieve a desired mapping. For example, the hyperbolic tangent function can be modified so that the mapping between the target rate and the error is as follows: target rate = a1 + tanh(a2error + a3), where a1, a2, and a3 are constants that depend on the maximum target rate and the three thresholds 250, 252, and 254. Specifically, a1 can be used to set the value of the maximum asymptote, a2 can be used to set the rate at which the sigmoid function gradually increases to the maximum asymptote, and a3 can be used to set the crossover point from the sigmoid function to pure exponential braking. In some embodiments, the sigmoid function may form the entire mapping such that no other segmentation fragments may exist.

[0179] In other embodiments, a smooth function other than a sigmoid function may be used to define the mapping between the target rate and the error so that the target rate of zero error disappears. In some such embodiments, the target rate may be constrained to be below a saturation target rate.

[0180] For the tilt angle controller 44, the SMC module 62 may receive the tilt angle error φ err , the tilt angle error φ err The bank angle φ of the aircraft 10 can be indicated to be different from the command bank angle φ cmd In some embodiments, the tilt angle error in, is an indication of the aircraft's bank angle φ, may be a direct output of the sensor signal processing and estimation block 42 and may therefore be obtained by direct measurement, indirect measurement or by estimation. For example, It can be obtained by processing the output of a three-axis inertial measurement unit, or by using one or more derived relationships between the tilt angle φ and other known (directly or indirectly) measured or otherwise estimated state variables. The SMC block 62 can process the tilt angle error φ err , and outputs the target rate of change of the bank angle φ of the aircraft 10 The target rate of change of the tilt angle φ can then be is input to the FL block 64. In some embodiments, the SMC block 62 can be configured to err The absolute value of the bank angle φ of the aircraft 10 is assigned a target rate of change accordingly. The value of the tilt angle error φ err Target rate of change The resulting mapping between can be continuous, thus achieving Figure 8A In other embodiments, the mapping may be differentiable, so that Figure 8A The tilt angle error φ of the logical structure in err Target rate of change The differentiable mapping between can be expressed as the following if-else conditional statement:

[0181] If|φ err |≤C φ,3 then

[0182]

[0183] Else if|φ err |≤C φ,2 then

[0184]

[0185] Else if|φ err |≤C φ,1 then

[0186]

[0187]

[0188] wherein, is the saturation angular rate of the tilt angle, k φ is a constant, C φ,1 is the first threshold, C φ,2 is the second threshold, C φ,3 is the third threshold, and the PID function PID(φ err ) is a linear combination of the tilt angle error φ err , the differential of the tilt angle error φ err , and the integral of the tilt angle error φ err (expressed as where t is the current time, and t0 < t can be some time selected as part of the adjustment of the PID function). k, C k, C φ,1 , C φ,2 , C φ,3 , t0 and the constants in the linear combination of PID(φ err ) can be selected to obtain the desired performance from the SMC block 62, and limit the maximum target change rate of the tilt angle φ to the desired maximum target angular rate As generally described above with reference to Figure 8C In some embodiments, and at least a part of the mapping between φ err can be a sigmoid function.

[0189] For the heading angle controller, the SMC module 62 can receive the heading angle error ψ err , which err can indicate the difference between the heading angle ψ of the aircraft 10 and the commanded heading angle ψ cmd . In some embodiments, the heading angle error wherein, is an indication of the aircraft heading angle ψ, can be the direct output of the sensor signal processing and estimation block 42, and thus can be obtained by direct measurement, indirect measurement, or by estimation. For example, can be obtained by processing and combining the outputs of a triaxial inertial measurement unit and a magnetic field sensor, or by using one or more derived relationships between the heading and other (direct or indirect) measured, or otherwise estimated, state variables. The SMC block 62 can process the heading angle error ψ err , and output the target change rate of the heading angle (ψ) of the aircraft 10. Then, the target change rate Provided to the FL block 64. In some embodiments, the SMC block 62 may test the heading angle error ψ against a series of conditions err by taking the absolute value and accordingly assigning a target rate of change such that the resulting mapping between the heading angle error ψ err and the target rate of change can be continuous, thereby implementing the Figure 8A logical structure shown in. In other embodiments, the mapping may be differentiable. For example, the differentiable mapping between the heading angle error ψ Figure 8A implementing the err logical structure and the target rate of change can be represented as the following if-else conditional statement:

[0190] If|ψ err |≤C ψ,3 then

[0191]

[0192] Else if|ψ err |≤C ψ,2 then

[0193]

[0194] Else if|ψ err |≤C ψ,1 then

[0195]

[0196]

[0197] where, is the saturated angular rate, C ψ,1 is the first threshold, C ψ,2 is the second threshold, C ψ,3 is the third threshold, and the PID function PID(ψ err ) is a linear combination of the heading angle error ψ err , the derivative of the heading angle error ψ err , and the integral of the heading angle error ψ err (denoted as where t is the current time and t0 < t can be some time selected as part of the tuning of the PID function). The saturated rate max can be specified according to the maximum allowable tilt angle φ such that where, Indicates the true airspeed obtained via measurement, estimation, or other means, and g is the value of the acceleration of gravity estimated based on altitude and geographic location, approximated as the value of the acceleration of gravity at the surface of the Earth, or determined in other ways. err ) in the linear combination can be equal to 1. You can choose PID (ψ err ) and the constants in the linear combination of φ max ,C ψ,1 ,C ψ,2 ,C ψ,3 and t0 to obtain the desired performance from the SMC block 62 and to limit the tilt angle φ to a specified maximum tilt angle φ max . As mentioned above Figure 8C Generally speaking, in some embodiments, and ψ err At least a portion of the mapping between may be a sigmoid function.

[0198] For the altitude controller 50 , the SMC module 62 may receive the altitude error h err , the height error h err It can indicate the altitude h of the aircraft 10 and the command altitude h cmd In some embodiments, the height error in, is an indication of the aircraft's altitude h, may be a direct output of the sensor signal processing and estimation block 42 and thus may be obtained by direct measurement, indirect measurement or estimation. For example, It may be obtained by processing and combining the output of a pressure altimeter, a global positioning system (GPS) receiver, and / or radar, or by using one or more derived relationships between the altitude h and other (directly or indirectly) measured or otherwise estimated state variables. The SMC block 62 may process the altitude error h err , and outputs the target rate of change of the altitude h of the aircraft 10 The target rate of change of the height h can then be Provided to the FL block 64. In some embodiments, the SMC block 62 can be configured to measure the height error h by testing the height error h against a range of conditions. err and assign the target rate of change accordingly The value of height error h err Target rate of change The resulting mapping between can be continuous, thus achieving Figure 8A In other embodiments, the mapping may be differentiable, for example, to achieve Figure 8A The height error h of the logical structure in err Target rate of change The differentiable mapping between can be represented as the following if-else conditional statement:

[0199] If |h err | ≤ C h,3 then

[0200]

[0201] Else if |h err | ≤ C h,2 then

[0202]

[0203] Else if |h err | ≤ C h,1 then

[0204]

[0205]

[0206] Wherein, is the saturated vertical velocity, C h,1 is the first threshold, C h,2 is the second threshold, C h,3 is the third threshold, and the PID function PID(h err ) is a linear combination of the height error h err , the differential of the error h err , and the integral of the height error h err (expressed as Wherein, t is the current time, and t0 < t can be some time selected as part of the adjustment of the PID function). The saturated vertical velocity can be specified according to the maximum allowable flight path angle (FPA) γ max as Wherein, indicates the true airspeed obtained through measurement, estimation, or otherwise. The proportional term in the linear combination of PID(h ) can be equal to 1. The constants in the linear combination of PID(h err ), as well as the values of γ err , C max , C h,1 , C h,2 , C h,3 and t0 can be selected to obtain the desired performance from the SMC block 62 and limit the maximum FPA to γ max . As generally described above in reference to Figure 8C in some embodiments, and h errAt least a portion of the mapping between may be a sigmoid function.

[0207] Figure 9 A schematic diagram of an exemplary FL block 64A for controlling the bank angle φ of an aircraft in flight is shown. The FL block 64A may receive a target rate of change of the bank angle φ from the SMC block 62. and can use inputs from the sensor signal processing and estimation block 42 and the target rate of change To generate the target body roll rate P c As mentioned above, the tilt angle rate The relationship between the body roll rate P can be expressed as

[0208] In the FL block 64A, inputs can be provided that eliminate possible nonlinear plants and at the same time specify the desired dynamics, i.e., the rates of the state variables. The body roll rate P can be considered as an input to the plant, where the tilt angle φ is the state variable. The plant input P is set equal to the FL input P given by the following equation: c The nonlinear object can be approximately eliminated and the tilt angle rate is forced to Follow target rate of change The values ​​of may be inputs from the sensor signal processing and estimation block 42, representing the aircraft pitch angle, bank angle, body pitch rate, and body yaw rate, respectively. The FL block 64A may receive the values ​​of and calculate The result is provided to the summing junction 74 which may also receive the target rate of change from the SMC module 62. The output 76 of the summing junction 74 may provide the output of the FL block 64A. The output of the FL block 64A may then be provided to the airframe roll rate controller 46 (e.g., Figure 4 ), the body roll rate controller 46 can then deflect the control surface 20. The body roll rate controller 46 can be part of a multi-axis controller 47 that can simultaneously allow control of the body roll rate, body pitch rate, and body yaw rate.

[0209] Figure 10 A schematic diagram of an exemplary FL block 64B for controlling the heading angle ψ of an aircraft in flight is shown. The FL block 64B may receive a target rate of change of the heading angle ψ of the aircraft 10 from the SMC block 62. and can use inputs from the sensor signal processing and estimation block 42 and the target rate of change To generate the command tilt angle φ cmd For an aircraft performing a coordinated turn, the heading rate The relationship between the tilt angle φ can be expressed as Among them, V T is the true airspeed, and g is the acceleration due to gravity. As previously described, in the FL block 64B, inputs can be provided that eliminate possible nonlinearities in the plant while specifying the desired dynamics, i.e., the rates of the state variables. The bank angle φ can be viewed as an input to the plant, where the heading angle ψ is the state variable. The input bank angle φ for the plant is set equal to the command bank angle φ given by the following equation: cmd The FL input can approximately offset the nonlinear object and force the heading angular rate Follow target rate of change The FL block 64B may receive an input from the sensor signal processing and estimation block 42 indicating the aircraft true airspeed. calculate The result is provided as the output of the FL block 64B. The output of the FL block 64B can then be provided to the tilt angle controller 44 (e.g. Figure 5 ), the pitch angle controller 44 can then deflect the control surface 20. Thus, the pitch angle controller 44 can be an internal controller of the heading angle controller 48 (see Figure 5 ).

[0210] Figure 11 A schematic diagram of an exemplary FL block 64C for controlling the altitude h of an aircraft in flight is shown. The FL block 64C may receive a target rate of change of the altitude h from the SMC block 62. and can use inputs from the sensor signal processing and estimation block 42 and the target rate of change To generate the target change ΔT of thrust c The following energy equation can be used to deduce that under trim conditions (trim thrust T trim ) altitude rate of an aircraft flying under The relationship between ΔT and Among them, V Tis the true airspeed, g is the acceleration due to gravity, and m is the mass of the aircraft 10. As previously described, in the FL block 64C, inputs can be provided that eliminate possible nonlinearities in the plant while simultaneously specifying the desired dynamics, i.e., the rates of the state variables. The target change in thrust, ΔT, can be considered as an input to the plant, with altitude, h, being the state variable. The input ΔT to this plant is set equal to the FL input ΔT given by the following equation: c Can approximately cancel nonlinear objects and force the height rate Follow target rate of change in, may be an input from the sensor signal processing and estimation block 42 indicating the aircraft true airspeed, and The FL block 64C may receive the actual air acceleration from the sensor signal processing and estimation block 42. And will Input to product node 84. Product node 84 can also be from The (true air acceleration) estimator 82 receives an input which may be an indication of the true air acceleration In various embodiments, The (real air acceleration) estimator 82 can obtain the Direct numerical differentiation By using one or more known relationships between air acceleration and other measured or otherwise estimated quantities provided by the sensor signal processing and estimation block 42; Filtering and then numerically differentiating the result; or by any other suitable method apparent to those skilled in the art. The output of the product node can be represents the rate of change of kinetic energy per unit mass of the aircraft 10. This output may be an input to a summing junction 86 which may have a gain by a value g, delivering the target rate of change The output of summing node 86 can be multiplied by And the result may be the output of the FL block 64C. The output of the FL block 64C may then be provided to the thrust controller 52 (see Figure 6 ), then, the thrust controller 52 may ultimately cause a change in the engine operating condition, such as a change in the amount of fuel supplied to the combustor of the gas turbine engine. The thrust controller 52 may receive status information from the sensor signal processing and estimation unit 42. For example, the thrust controller 52 may receive a trim condition thrust T trim In some embodiments, the altitude controller 50 may be configured to adjust the target change in thrust ΔT based on the target change in thrust. c To calculate the TLA value, and feed the TLA value to the thrust controller 52.

[0211] Figure 12 A schematic representation of an exemplary bank angle control system 212 is shown. System 212 may have elements previously described above with respect to other systems, and therefore like reference numerals are used to identify like elements. Bank angle SMC block 62A may generate a target rate of change of bank angle φ and will Provided to the tilt angle FL block 64A, thus implementing the formula The FL block 64A can generate a target body roll rate P c , which can then be input into the body roll rate controller 46. The body roll rate controller 46 can generate one or more commands δ cmd , to cause the desired deflection of the flight control surface 20 via the actuator 34. The SMC block 62A and the FL block 64A can both receive inputs from the sensor signal processing and estimation block 42. The body roll rate controller 46 can include an aerodynamic model of the aircraft 10. The body roll rate controller 46 can utilize nonlinear dynamic inversion control techniques and / or incremental nonlinear dynamic inversion (INDI). INDI (sometimes referred to as "improved nonlinear dynamic inversion" or "modified nonlinear dynamic inversion") is a control technique that is based on calculating the incremental change in the control input required to steer the aircraft to a desired state. INDI is based on the assumption that for small time increments, the system's response to the control input is greater than its response to the state change. This assumption allows the control input increment to be calculated only from the system's input-dependent dynamics, while ignoring the system's state-dependent dynamics. The body roll rate controller 46 can also be enhanced with pseudo control limits (PCH) that can reduce the size of the command signal to a level that can be achieved by a saturated controller. State variables Can be a pseudo measurement.

[0212] Figure 13 A schematic representation of an exemplary heading angle control system 214 is shown. The system 214 may have elements previously described above with respect to other systems, and therefore like reference numerals are used to identify like elements. The heading angle SMC block 62B may generate a target rate of change of the heading angle (ψ) and provides it to the heading angle FL block 64B, thus realizing the formula The FL block 64B can generate a commanded tilt angle φ cmd , then the command can be tilted by an angle φ cmd The target body roll rate P is input to the tilt angle controller 44. The tilt angle controller 44 can generate a target body roll rate P c, then the target body roll rate P c The input is to the body roll rate controller 46. The body roll rate controller 46 generates one or more commands δ cmd , to cause a desired deflection of the flight control surface 20 via the actuator 34 . Both the SMC block 62B and the FL block 64B may receive input from the sensor signal processing and estimation block 42 .

[0213] Figure 14 A schematic representation of an exemplary altitude control system 216 is shown. System 216 may have elements previously described above with respect to other systems, and thus like reference numerals are used to identify like elements. Altitude SMC block 62C may generate a target rate of change of altitude h And will Provided to the height FL block 64C, thereby implementing the formula The FL block 64C may generate the command dT c , the command dT c is the target change in thrust. Then the command dT c Add to trim condition thrust T trim , and the resulting total thrust is provided to the thruster 210. The thruster 210 can change the engine settings to change the thrust. Both the SMC block 62C and the FL block 64C can receive input from the sensor signal processing and estimation block 42.

[0214] Figure 15 is a flow chart illustrating a method 100 for controlling the bank angle φ of an aircraft during flight. The method 100 may be performed using the system 12 described herein or using other systems. For example, the machine readable instructions 32 (see Figure 2 ) can be configured to cause the computer 24 to perform at least a portion of the method 100. It should be understood that aspects of the method 100 can be combined with aspects of other methods described herein. In various embodiments, the method 100 may include:

[0215] Receive the commanded bank angle φ of the aircraft 10 cmd (See Box 102);

[0216] Calculate the tilt angle error φ err , the tilt angle error φ err The bank angle φ of the aircraft 10 is indicated and the command bank angle φ cmd the difference between (see box 104);

[0217] The target rate of change of the bank angle φ of the aircraft 10 is calculated using a sliding mode control technique. The input to the sliding mode control technique includes the tilt angle error φ err (See Box 106);

[0218] The target body roll rate P of the aircraft 10 is calculated using the FL control technique. c The input to the FL control technique includes the target rate of change of the bank angle φ of the aircraft 10 (See Box 108); and

[0219] Use target body roll rate P during flight c To control one or more actuators 34 of the aircraft 10 .

[0220] FL control techniques may include calculating a target body roll rate P for the aircraft 10 using the inverse of the relationship between the bank angle φ of the aircraft 10 and one or more body angular rates of the aircraft 10 . c .

[0221] FL control techniques may include calculating a target airframe roll rate using the following formula:

[0222]

[0223] in, represents the value of the pitch angle of the aircraft 10, represents the value of the bank angle φ of the aircraft 10, a value representing the body pitch rate of the aircraft 10, and A value indicating the body yaw rate of the aircraft 10 .

[0224] Sliding mode control techniques can include the following: err , the first threshold C φ,1 , the second threshold C φ,2 and the third threshold C φ,3 , generating the target rate of change of the tilt angle φ When the tilt angle error φ err The absolute value of the φ,1 When the target change rate of the tilt angle φ is is chosen to be substantially equal to the tilt angle saturation rate

[0225] When the tilt angle error φ err The absolute value of is less than the first threshold C φ,1 and is greater than the second threshold C φ,2 The target rate of change of the tilt angle φ can be calculated using the following formula:

[0226]

[0227] Among them, sign(φ err ) is the tilt angle error φ errThe symbolic function of k φ Represents a parameter (such as a constant).

[0228] When the tilt angle error φ err The absolute value of is less than the third threshold C φ,3 When the tilt angle error φ err The target rate of change of the tilt angle φ is selected by using the proportional-integral-derivative control function

[0229] When the tilt angle error φ err The absolute value of is less than the second threshold C φ,2 and is greater than the third threshold C φ,3 When the target change rate of the tilt angle φ is Selected to be equal to the tilt angle error φ err Proportional.

[0230] When the tilt angle error φ err The absolute value of is less than the second threshold C φ,2 and is greater than the third threshold C φ,3 , the following formula can be used: To calculate the target rate of change of the tilt angle φ

[0231] Figure 16 is a flow chart illustrating a method 120 for controlling the heading angle ψ of an aircraft during flight. The method 120 may be performed using the system 12 described herein or using other systems. For example, the machine readable instructions 32 (see Figure 2 ) can be configured to cause the computer 24 to perform at least a portion of the method 120. It should be understood that aspects of the method 120 can be combined with aspects of other methods described herein. In various embodiments, the method 120 may include:

[0232] Receive a commanded heading angle ψ for the aircraft 10 cmd (See Box 122);

[0233] Calculate the heading angle error ψ err , the heading angle error ψ err Indicates the aircraft's heading angle ψ and the commanded heading angle ψ cmd the difference between (see box 124);

[0234] The target rate of change of the heading angle ψ of the aircraft 10 is calculated using a sliding mode control technique. The input to the sliding mode control technique includes the heading angle error ψ err (See Box 126);

[0235] The FL control technique is used to calculate the commanded bank angle φ of the aircraft 10 cmd, the input to the FL control technique includes the target rate of change of the heading angle ψ of the aircraft 10 (See Box 128); and

[0236] During flight, the commanded bank angle φ is used cmd to control one or more actuators 34 of the aircraft (see block 130 ).

[0237] FL control techniques may include calculating a commanded bank angle φ using the inverse of the relationship between the heading angle ψ of the aircraft 10 and the bank angle φ of the aircraft 10 . cmd .

[0238] FL control techniques may include using the following formula: To calculate the command tilt angle φ cmd ,in, represents the true airspeed of the aircraft 10 , and g represents the acceleration due to gravity.

[0239] The sliding mode control technique may include the following steps based on the heading angle error ψ err , the first threshold C ψ,1 , the second threshold C ψ,2 and the third threshold C ψ,3 , generating the target rate of change of the heading angle ψ So that when the heading angle error ψ err The absolute value of the ψ,1 When the target change rate of the heading angle ψ is is chosen to be substantially equal to the heading angle saturation rate

[0240] When the heading angle error ψ err The absolute value of is less than the first threshold C ψ,1 and is greater than the second threshold C ψ,2 , the following formula can be used: To calculate the target rate of change of the heading angle ψ Among them, sign(ψ err ) is the heading angle error ψ err The sign function of .

[0241] When the heading angle error ψ err The absolute value of is less than the third threshold C ψ,3 When the heading angle error ψ err The target rate of change of the heading angle ψ is selected using the proportional-integral-derivative control function

[0242] When the heading angle error ψ err The absolute value of is less than the second threshold C ψ,2 and is greater than the third threshold C ψ,3 When the target change rate of the heading angle ψ is is chosen to be substantially equal to the heading angle error ψ err Proportional.

[0243] When the heading angle error ψ err The absolute value of is less than the second threshold C ψ,2 and is greater than the third threshold C ψ,3 When , the following formula can be used: To calculate the target rate of change of the heading angle ψ

[0244] When the heading angle error ψ err The absolute value of is less than the third threshold C ψ,3 When , the proportional term in the proportional-integral-derivative control function can be basically equal to

[0245] In some embodiments, methods 100 and 120 may be combined. For example, method 100 may be appended to method 120. Thus, the FL control technique of method 120 may be a first FL control technique; the sliding mode control technique of method 120 may be a first sliding mode control technique; and the commanded bank angle φ is used during flight. cmd The one or more actuators 34 used to control the aircraft 10 may include:

[0246] Calculate the tilt angle error φ err , the tilt angle error φ err The bank angle φ of the aircraft 10 is indicated and the command bank angle φ cmd The difference between

[0247] The target rate of change of the bank angle φ of the aircraft 10 is calculated using the second sliding mode control technique. The input to the second sliding mode control technique includes the tilt angle error φ err ;

[0248] The target body roll rate P of the aircraft 10 is calculated using the second FL control technique. c The input to the second FL control technique includes the target rate of change of the bank angle φ of the aircraft 10 as well as

[0249] Use target body roll rate P during flight c to control the one or more actuators 34 of the aircraft 10 .

[0250] Figure 17 is a flow chart illustrating a method 140 for controlling the altitude h of an aircraft during flight. The method 140 may be performed using the system 12 described herein or using other systems. For example, the machine readable instructions 32 (see Figure 2) can be configured to cause the computer 24 to perform at least a portion of the method 140. It should be understood that aspects of the method 140 can be combined with aspects of other methods described herein. In various embodiments, the method 140 may include:

[0251] Receive a commanded altitude h for the aircraft 10 cmd (See Box 142);

[0252] Calculate the height error h err , the height error h err Indicates the altitude h of the aircraft 10 and the commanded altitude h cmd the difference between (see box 144);

[0253] The target rate of change of the altitude h of the aircraft 10 is calculated using a sliding mode control technique. The input to the sliding mode control technique includes the height error h err (See Box 146);

[0254] A value indicative of a target change in thrust for the aircraft 10 is calculated using a FL control technique, the inputs to which include a target rate of change of the altitude h of the aircraft 10 as well as

[0255] The values ​​are used to control one or more actuators 34 of the aircraft 10 during flight.

[0256] The value may be a target change in TLA for the aircraft 10 .

[0257] FL control techniques may include calculating the value using the inverse of the relationship between the altitude h of the aircraft 10 and the airspeed of the aircraft 10 .

[0258] FL control techniques can include determining the aircraft's true airspeed Rate of change and using the following formula: To calculate the target change in thrust ΔT c , where g represents the value of gravitational acceleration and m represents the mass of the aircraft.

[0259] Sliding mode control techniques can include the following: err , the first threshold C h,1 , the second threshold C h,2 and the third threshold C h,3 , generating the target rate of change of height h So that when the height error h err The absolute value of the h,1 When the target change rate of the aircraft 10's altitude h is is chosen to be substantially equal to the high saturation rate

[0260] When the height error h err The absolute value of is less than the first threshold C h,1 and is greater than the second threshold C h,2 , the following formula can be used: To calculate the target rate of change of height h Among them, sign(h err ) is the height error h err The sign function of .

[0261] When the height error h err The absolute value of is less than the third threshold C h,3 When the height error h err The proportional-integral-derivative control function is used to select the target rate of change of the height h

[0262] When the height error h err The absolute value of is less than the second threshold C h,2 and is greater than the third threshold C h,3 When the target change rate of height h is Choose the height error h err Proportional.

[0263] When the height error h err The absolute value of is less than the second threshold C h,2 and is greater than the third threshold C h,3 , the following formula can be used: To calculate the target rate of change of height h

[0264] When the height error h err The absolute value of is less than the third threshold C h,3 When , the proportional term in the proportional-integral-derivative control function can be basically equal to

[0265] The above description is merely exemplary, and those skilled in the relevant art will recognize that changes may be made to the described embodiments without departing from the scope of the disclosed invention. The present disclosure can be embodied in other specific forms without departing from the subject matter of the claims. The present disclosure is intended to cover and encompass all suitable changes in technology. Modifications that fall within the scope of the present invention will be apparent to those skilled in the art upon review of the present disclosure, and such modifications are intended to fall within the appended claims. Moreover, the scope of the claims should not be limited by the preferred embodiments set forth in the examples, but should be given the broadest interpretation consistent with the entire specification.

Claims

1. A method for controlling the bank angle (φ) of an aircraft during flight, the method comprising: Receive the commanded bank angle (φ) for the aircraft cmd ); Calculate the tilt angle error (φ err ), the tilt angle error (φ err ) indicates the aircraft's bank angle (φ) and the command bank angle (φ cmd ) Use sliding mode control techniques to calculate the target rate of change of the aircraft's bank angle (φ) The input to the sliding mode control technique includes the tilt angle error (φ err ); The target body roll rate (P c ), the inputs to the FL control technique include the target rate of change of the aircraft's bank angle (φ) as well as During flight, the target body roll rate (P c ) to control one or more actuators of the aircraft; The FL control technique includes calculating a target body roll rate (P) for the aircraft using the inverse relationship between the aircraft's bank angle (φ) and one or more body angular rates of the aircraft. c ).

2. The method according to claim 1, wherein The FL control technique includes calculating a target body roll rate (P) for the aircraft using the inverse of the relationship between the aircraft's bank angle (φ) and one or more body angular rates of the aircraft. c ).

3. The method according to claim 1, wherein The FL control technique includes calculating the target body roll rate using the following formula: in, represents the value of the pitch angle of the aircraft, represents the value of the bank angle (φ) of the aircraft, represents a value of the aircraft's body pitch rate, and A value representing the body yaw rate of the aircraft.

4. The method according to any one of claims 1 to 3, wherein: The sliding mode control technology includes the following steps: err ), the first threshold (C φ，1 ), the second threshold (C φ，2 ) and the third threshold (C φ，3 ), generating the target rate of change of the tilt angle (φ) When the tilt angle error (φ err ) is greater than the first threshold value (C φ，1 ), the target change rate of the tilt angle (φ) is chosen to be substantially equal to the tilt angle saturation rate 5. The method according to claim 4, wherein When the tilt angle error (φ err ) is less than the first threshold value (C φ，1 ) and is greater than the second threshold (C φ，2 ), use the following formula: To calculate the target rate of change of the tilt angle (φ) Among them, sign(φ err ) is the tilt angle error (φ err ) is a sign function, and k φ Indicates a parameter.

6. The method according to claim 4 or 5, wherein: When the tilt angle error (φ err ) is less than the third threshold value (C φ，3 ), based on the tilt angle error (φ err ) to select the target rate of change of the tilt angle (φ) using the proportional-integral-derivative control function 7. The method according to any one of claims 4 to 6, wherein: When the tilt angle error (φ err ) is less than the second threshold value (C φ，2 ) and is greater than the third threshold (C φ，3 ), the target change rate of the tilt angle (φ) is selected to be equal to the tilt angle error (φ err ) are proportional.

8. The method according to any one of claims 4 to 7, wherein: When the tilt angle error (φ err ) is less than the second threshold value (C φ，2 ) and is greater than the third threshold (C φ，3 ), use the following formula: To calculate the target rate of change of the tilt angle (φ) Among them, k φ Indicates a parameter.

9. The method according to any one of claims 1 to 3, wherein: The sliding mode control technique includes using a sigmoid function as the target rate of change of the tilt angle (φ) The tilt angle error (φ err ) between them.

10. The method according to any one of claims 1 to 9, wherein The aircraft is a blended wing-body aircraft.

11. A computer program product for implementing a bank angle control function of an aircraft during flight, the computer program product comprising a non-transitory machine-readable storage medium having program code embodied therein, the program code being readable / executable by a computer, a processor, or a logic circuit to perform the method according to any one of claims 1 to 10.

12. A system for controlling the bank angle (φ) of an aircraft during flight, the system comprising: One or more computers operatively coupled to receive a commanded bank angle (φ) for the aircraft. cmd ), the one or more computers being configured to: Calculate the tilt angle error (φ err ), the tilt angle error (φ err ) indicates the aircraft's bank angle (φ) and the command bank angle (φ cmd ) Use sliding mode control techniques to calculate the target rate of change of the aircraft's bank angle (φ) The input to the sliding mode control technique includes the tilt angle error (φ err ); The target body roll rate (P c ), the inputs to the FL control technique include the target rate of change of the aircraft's bank angle (φ) as well as During flight, the target body roll rate (P c ) to control one or more actuators of the aircraft; The FL control technique includes calculating a target body roll rate (P) for the aircraft using the inverse relationship between the aircraft's bank angle (φ) and one or more body angular rates of the aircraft. c ).

13. The system according to claim 12, wherein: The FL control technique includes calculating a target body roll rate (P) for the aircraft using the inverse of the relationship between the aircraft's bank angle (φ) and one or more body angular rates of the aircraft. c ).

14. The system according to claim 12, wherein: The FL control technique includes calculating the target body roll rate using the following formula: in, represents the value of the pitch angle of the aircraft, represents the value of the bank angle (φ) of the aircraft, represents a value of the aircraft's body pitch rate, and A value representing the body yaw rate of the aircraft.

15. The system according to any one of claims 12 to 14, wherein: The sliding mode control technology includes the following steps: err ), the first threshold (C φ，1 ), the second threshold (C φ，2 ) and the third threshold (C φ，3 ), generating the target rate of change of the tilt angle (φ) When the tilt angle error (φ err ) is greater than the first threshold value (C φ，1 ), the target change rate of the tilt angle (φ) is chosen to be substantially equal to the tilt angle saturation rate 16. The system according to claim 15, wherein: When the tilt angle error (φ err ) is less than the first threshold value (C φ，1 ) and is greater than the second threshold (C φ，2 ), use the following formula: To calculate the target rate of change of the tilt angle (φ) Among them, φ err represents the tilt angle error, sign(φ err ) is the tilt angle error (φ err ) is a sign function, and k φ Indicates a parameter.

17. A system according to claim 15 or claim 16, wherein: When the tilt angle error (φ err ) is less than the third threshold value (C φ，3 ), based on the tilt angle error (φ err ) to select the target rate of change of the tilt angle (φ) using the proportional-integral-derivative control function 18. A system according to any one of claims 15 to 17, wherein When the tilt angle error (φ err ) is less than the second threshold value (C φ，2 ) and is greater than the third threshold (C φ，3 ), the target change rate of the tilt angle (φ) is selected to be substantially equal to the tilt angle error (φ err ) are proportional.

19. The system according to any one of claims 15 to 18, wherein: When the tilt angle error (φ err ) is less than the second threshold value (C φ，2 ) and is greater than the third threshold (C φ，3 ), according to the following formula: To calculate the target rate of change of the tilt angle (φ) Among them, k φ Indicates a parameter.

20. The system according to any one of claims 12 to 14, wherein The sliding mode control technique includes using a sigmoid function as the target rate of change of the tilt angle (φ) The tilt angle error (φ err ) between them.

21. An aircraft comprising a system according to any one of claims 12 to 20.

22. A blended wing body aircraft comprising a system according to any one of claims 12 to 20.

23. A method for controlling the heading angle (ψ) of an aircraft during flight, the method comprising: Receive the command heading angle (ψ cmd ); Calculate the heading angle error (ψ err ), the heading angle error (ψ err ) indicates the heading angle of the aircraft (ψ) and the command heading angle (ψ cmd ) Use sliding mode control techniques to calculate the target rate of change of the aircraft's heading angle (ψ) The input to the sliding mode control technique includes the heading angle error (ψ err ); The feedback linearization (FL) control technique is used to calculate the commanded bank angle (φ) of the aircraft. cmd ), the inputs to the FL control technique include the target rate of change of the aircraft's heading angle (ψ) as well as During flight, the commanded tilt angle (φ cmd ) to control one or more actuators of the aircraft; The FL control technique includes calculating the commanded bank angle (φ) using the inverse of the relationship between the aircraft's heading angle (ψ) and the aircraft's bank angle (φ). cmd ).

24. The method according to claim 23, wherein The FL control technique involves calculating the commanded bank angle (φ) using the inverse of the relationship between the aircraft's heading angle (ψ) and the aircraft's bank angle (φ). cmd ).

25. The method according to claim 23, wherein The FL control technique involves using the following formula: To calculate the command tilt angle (φ cmd ),in, represents the true airspeed of the aircraft, and g represents the acceleration due to gravity.

26. The method according to any one of claims 23 to 25, wherein The sliding mode control technology includes the following steps: err ), the first threshold (C ψ，1 ), the second threshold (C ψ，2 ) and the third threshold (C ψ，3 ), generating the target rate of change of the heading angle (ψ) So that when the heading angle error (ψ err ) is greater than the first threshold value (C ψ，1 ), the target change rate of the heading angle (ψ) is chosen to be substantially equal to the heading angle saturation rate 27. The method according to claim 26, wherein When the heading angle error (ψ err ) is less than the first threshold value (C ψ，1 ) and is greater than the second threshold (C ψ，2 ), use the following formula: To calculate the target rate of change of the heading angle (ψ) Among them, sign(ψ err ) is the heading angle error (ψ err ) is the sign function of .

28. The method according to claim 26 or 27, wherein When the heading angle error (ψ err ) is less than the third threshold value (C ψ，3 ), based on the heading angle error (ψ err ) to select the target rate of change of the heading angle (ψ) using the proportional-integral-derivative control function 29. The method according to claim 28, wherein When the heading angle error (ψ err ) is less than the third threshold value (C ψ，3 ), the proportional term in the proportional-integral-differential control function is substantially equal to 30. The method according to any one of claims 26 to 29, wherein When the heading angle error (ψ err ) is less than the second threshold value (C ψ，2 ) and is greater than the third threshold (C ψ，3 ), the target change rate of the heading angle (ψ) is selected to be substantially equal to the heading angle error (ψ err ) are proportional.

31. The method according to any one of claims 26 to 30, wherein When the heading angle error (ψ err ) is less than the second threshold value (C ψ，2 ) and is greater than the third threshold (C ψ，3 ), according to the following formula: To calculate the target rate of change of the heading angle (ψ) 32. The method according to any one of claims 23 to 25, wherein: The sliding mode control technique includes using a sigmoid function as the target rate of change of the heading angle (ψ) The heading angle error (ψ err ) between them.

33. The method according to any one of claims 23 to 32, wherein: The FL control technique is a first FL control technique; The sliding mode control technology is a first sliding mode control technology; as well as During flight, the commanded tilt angle (φ cmd ) to control the one or more actuators of the aircraft include: Calculate the tilt angle error (φ cmd ), the tilt angle error (φ cmd ) indicates the aircraft's bank angle (φ) and the command bank angle (φ cmd ) The second sliding mode control technique is used to calculate the target rate of change of the aircraft's bank angle (φ) The input to the second sliding mode control technique includes the tilt angle error (φ err ); The second FL control technique is used to calculate the target body roll rate (P c ), the input to the second FL control technique includes the target rate of change of the aircraft's bank angle (φ) as well as During flight, the target body roll rate (P c ) to control the one or more actuators of the aircraft.

34. The method according to any one of claims 23 to 33, wherein: The aircraft is a blended wing-body aircraft.

35. A computer program product for implementing a heading angle control function of an aircraft during flight, the computer program product comprising a non-transitory machine-readable storage medium having program code embodied therein, the program code being readable / executable by a computer, a processor, or a logic circuit to perform the method according to any one of claims 23 to 34.

36. A system for controlling the heading angle (ψ) of an aircraft during flight, the system comprising: One or more computers operatively coupled to receive a commanded heading angle (ψ) for the aircraft. cmd ), the one or more computers being configured to: Calculate the heading angle error (ψ err ), the heading angle error (ψ err ) indicates the heading angle of the aircraft (ψ) and the command heading angle (ψ cmd ) Use sliding mode control techniques to calculate the target rate of change of the aircraft's heading angle (ψ) The input to the sliding mode control technique includes the heading angle error (ψ err ); The feedback linearization (FL) control technique is used to calculate the commanded bank angle (φ) of the aircraft. cmd ), the inputs to the FL control technique include the target rate of change of the aircraft's heading angle (ψ) as well as During flight, the commanded tilt angle (φ cmd ) to control one or more actuators of the aircraft; The FL control technique includes calculating the commanded bank angle (φ) using the inverse of the relationship between the aircraft's heading angle (ψ) and the aircraft's bank angle (φ). cmd ).

37. The system of claim 36, wherein: The FL control technique involves calculating the commanded bank angle (φ) using the inverse of the relationship between the aircraft's heading angle (ψ) and the aircraft's bank angle (φ). cmd ).

38. The system of claim 36, wherein: The FL control technique involves using the following formula: To calculate the command tilt angle (φ cmd ),in, represents the true airspeed of the aircraft, and g represents the acceleration due to gravity.

39. A system according to any one of claims 36 to 38, wherein The sliding mode control technology includes the following steps: err ), the first threshold (C ψ，1 ), the second threshold (C ψ，2 ) and the third threshold (C ψ，3 ), generating the target rate of change of the heading angle (ψ) So that when the heading angle error (ψ err ) is greater than the first threshold value (C ψ，1 ), the target change rate of the heading angle (ψ) is chosen to be substantially equal to the heading angle saturation rate 40. The system of claim 39, wherein: When the heading angle error (ψ err ) is less than the first threshold value (C ψ，1 ) and is greater than the second threshold (C ψ，2 ), use the following formula: To calculate the target rate of change of the heading angle (ψ) Among them, sign(ψ err ) is the heading angle error (ψ err ) is the sign function of .

41. The system of claim 39 or 40, wherein: When the heading angle error (ψ err ) is less than the third threshold value (C ψ，3 ), based on the heading angle error (ψ err ) to select the target rate of change of the heading angle (ψ) using the proportional-integral-derivative control function 42. The system of claim 41, wherein: When the heading angle error (ψ err ) is less than the third threshold value (C ψ，3 ), the proportional term in the proportional-integral-differential control function is substantially equal to 43. A system according to any one of claims 39 to 42, wherein: When the heading angle error (ψ err ) is less than the second threshold value (C ψ，2 ) and is greater than the third threshold (C ψ，3 ), the target change rate of the heading angle (ψ) is selected to be substantially equal to the heading angle error (ψ err ) are proportional.

44. A system according to any one of claims 39 to 43, wherein When the heading angle error (ψ err ) is less than the second threshold value (C ψ，2 ) and is greater than the third threshold (C ψ，3 ), according to the following formula: To calculate the target rate of change of the heading angle (ψ) 45. The system of any one of claims 36 to 38, wherein: The sliding mode control technique includes using a sigmoid function as the target rate of change of the heading angle (ψ) The heading angle error (ψ err ) between them.

46. ​​A system according to any one of claims 36 to 45, wherein: The FL control technique is a first FL control technique; The sliding mode control technology is a first sliding mode control technology; as well as During flight, the commanded tilt angle (φ cmd ) to control the one or more actuators of the aircraft include: Calculate the tilt angle error (φ cmd ), the tilt angle error (φ cmd ) indicates the aircraft's bank angle (φ) and the command bank angle (φ cmd ) The second sliding mode control technique is used to calculate the target rate of change of the aircraft's bank angle (φ) The input to the second sliding mode control technique includes the tilt angle error (φ err ); The second FL control technique is used to calculate the target body roll rate (P c ), the inputs to the FL control technique include the target rate of change of the aircraft's bank angle (φ) as well as During flight, the target body roll rate (P c ) to control the one or more actuators of the aircraft.

47. An aircraft comprising a system according to any one of claims 36 to 46.

48. A blended wing-body aircraft comprising the system of any one of claims 36 to 46.

49. A method for controlling the altitude (h) of an aircraft during flight, the method comprising: Receive the commanded altitude (h cmd ); Calculate the height error (h err ), the height error (h err ) indicates the altitude of the aircraft (h) and the command altitude (h cmd ) Use sliding mode control techniques to calculate the target rate of change of the aircraft's altitude (h) The input to the sliding mode control technique includes the height error (h err ); A value representing a target change in thrust for the aircraft is calculated using a feedback linearization (FL) control technique, the input to which includes a target rate of change of the aircraft's altitude (h) as well as using the value to control one or more actuators of the aircraft during flight; Wherein the FL control technique includes calculating the value using the inverse of the relationship between the aircraft's altitude (h) and the aircraft's airspeed.

50. The method of claim 49, wherein The value is the target change in thrust lever angle for the aircraft.

51. The method according to claim 49 or 50, wherein The FL control technique includes calculating the value using the inverse of the relationship between the aircraft's altitude (h) and the aircraft's airspeed.

52. The method of claim 49, wherein The FL control technology includes: Determine the true airspeed of the aircraft Rate of change as well as Use the following formula: To calculate the target change in thrust (ΔT c ), where g represents the value of gravitational acceleration and m represents the mass of the aircraft.

53. The method according to any one of claims 49 to 52, wherein: The sliding mode control technology includes the following steps: err ), the first threshold (C h，1 ), the second threshold (C h，2 ) and the third threshold (C h，3 ), generating the target rate of change of the height (h) So that when the height error (h err ) is greater than the first threshold value (C h，1 ), the target change rate of the height (h) is chosen to be substantially equal to the high saturation rate 54. The method of claim 53, wherein: When the height error (h err ) is less than the first threshold value (C h，1 ) and is greater than the second threshold (C h，2 ), use the following formula: To calculate the target rate of change of the height (h) Among them, sign(h err ) is the height error (h err ) is the sign function of .

55. The method according to claim 53 or 54, wherein When the height error (h err ) is less than the third threshold value (C h，3 ), based on the height error (h err ) proportional-integral-derivative control function to select the target rate of change of the height (h) 56. The method of claim 55, wherein: When the height error (h err ) is less than the third threshold value (C h，3 ), the proportional term in the proportional-integral-differential control function is substantially equal to 57. The method according to any one of claims 53 to 55, wherein When the height error (h err ) is less than the second threshold value (C h，2 ) and is greater than the third threshold (C h，3 ), the target change rate of the height (h) is selected to be equal to the height error (h err ) are proportional.

58. The method according to any one of claims 53 to 56, wherein When the height error (h err ) is less than the second threshold value (C h，2 ) and is greater than the third threshold (C h，3 ), use the following formula: To calculate the target rate of change of the height (h) 59. The method according to any one of claims 49 to 51, wherein The sliding mode control technique includes using a sigmoid function as the target rate of change of the height (h) The height error (h err ) between them.

60. The method according to any one of claims 49 to 59, wherein The aircraft is a blended wing-body aircraft.

61. A computer program product for implementing an altitude control function for an aircraft during flight, the computer program product comprising a non-transitory machine-readable storage medium having program code embodied therein, the program code being readable / executable by a computer, a processor, or a logic circuit to perform the method according to any one of claims 49 to 60.

62. A system for controlling the altitude (h) of an aircraft during flight, the system comprising: One or more computers operatively coupled to indicate a commanded altitude (h cmd ), the one or more computers being configured to: Calculate the height error (h err ), the height error (h err ) indicates the altitude (h) and command altitude (h) of the aircraft cmd ) Calculate the target rate of change of the aircraft's altitude (h) using sliding mode control techniques The input to the sliding mode control technique includes the height error (h err ); A value representing a target change in thrust for the aircraft is calculated using a feedback linearization control technique, the input to the FL control technique including the target rate of change of the aircraft's altitude (h) as well as using the value to control one or more actuators of the aircraft during flight; Wherein the FL control technique includes calculating the value using the inverse of the relationship between the aircraft's altitude (h) and the aircraft's airspeed.

63. The system of claim 62, wherein: The value is the target change in thrust lever angle for the aircraft.

64. The system of claim 62 or 63, wherein: The FL control technique includes calculating the value using the inverse of the relationship between the aircraft's altitude (h) and the aircraft's airspeed.

65. The system of claim 62, wherein: The FL control technology includes: Determine the true airspeed of the aircraft Rate of change as well as Use the following formula: To calculate the target change in thrust rod angle (ΔT c ), where g represents the value of gravitational acceleration and m represents the mass of the aircraft.

66. A system according to any one of claims 62 to 65, wherein The sliding mode control technology includes the following steps: err ), the first threshold (C h，1 ), the second threshold (C h，2 ) and the third threshold (C h，3 ), generating the target rate of change of the height (h) So that when the height error (h err ) is greater than the first threshold value (C h，1 ), the target change rate of the height (h) is chosen to be substantially equal to the high saturation rate 67. The system of claim 66, wherein: When the height error (h err ) is less than the first threshold value (C h，1 ) and is greater than the second threshold (C h，2 ), use the following formula: To calculate the target rate of change of the height (h) Among them, sign(h err ) is the height error (h err ) is the sign function of .

68. A system according to claim 66 or claim 67, wherein When the height error (h err ) is less than the third threshold value (C h，3 ), based on the height error (h err ) proportional-integral-derivative control function to select the target rate of change of the height (h) 69. The system of claim 68, wherein When the height error (h err ) is less than the third threshold value (C h，3 ), the proportional term in the proportional-integral-differential control function is substantially equal to 70. The system of any one of claims 66 to 69, wherein: When the height error (h err ) is less than the second threshold value (C h，2 ) and is greater than the third threshold (C h，3 ), the target change rate of the height (h) is selected to be equal to the height error (h err ) are proportional.

71. The system of any one of claims 66 to 69, wherein: When the height error (h err ) is less than the second threshold value (C h，2 ) and is greater than the third threshold (C h，3 ), use the following formula: To calculate the target rate of change of the height (h) 72. The system of any one of claims 62 to 64, wherein: The sliding mode control technique includes using a sigmoid function as the target rate of change of the height (h) The height error (h err ) between them.

73. An aircraft comprising a system according to any one of claims 62 to 72.

74. A blended wing body aircraft comprising a system according to any one of claims 62 to 72.

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