Aircraft control systems and methods using sliding mode control and feedback linearization

Patent metadata

Field	Value
Publication number	US-12386362-B2
Application number	US-202418660437-A
Country	US
Kind code	B2
Filing date	May 10, 2024
Priority date	Sep 30, 2019
Publication date	Aug 12, 2025
Grant date	Aug 12, 2025

Abstract

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

First claim

What is claimed is:

1. A method for controlling a heading angle (ψ) of an aircraft during flight, the method comprising:
receiving a commanded heading angle (ψ cmd ) for the aircraft;
computing a heading angle error (ψ_err) indicative of a difference between the heading angle (ψ) of the aircraft and the commanded heading angle (ψ cmd );
computing a target rate of change ({dot over (ψ)} des ) for the heading angle (ψ) of the aircraft using a sliding mode control technique, an input of the sliding mode control technique including the heading angle error (ψ_err);
computing a commanded bank angle (ϕ cmd ) for the aircraft using a feedback linearization (FL) control technique, an input to the FL control technique including the target rate of change ({dot over (ψ)} des ) for the heading angle (ψ) of the aircraft; and
using the commanded bank angle (ϕ cmd ) to control one or more actuators of the aircraft during flight.

2. The method of claim 1 , wherein the FL control technique includes using an inversion of a relationship between the heading angle (ψ) of the aircraft and a bank angle (ϕ) of the aircraft to calculate the commanded bank angle (ϕ cmd ).

3. The method of claim 1 , wherein the FL control technique includes computing the commanded bank angle (ϕ cmd ) using the following formula:

ϕ
cmd

=

arctan
⁢

(

ψ
.

des

⁢

V
^

T

g

)

,

where {circumflex over (V)} T is indicative of a true air speed of the aircraft, and g denotes gravitational acceleration.

4. The method of claim 1 , wherein the sliding mode control technique includes generating the target rate of change ({dot over (ψ)} des ) of the heading angle (ψ), as a function of the heading angle error (ψ_err), a first threshold (C ψ,1 )), a second threshold (C ψ,2 ), and a third threshold (C ψ,3 ) such that the target rate of change ({dot over (ψ)} des ) of the heading angle (ψ), when an absolute value of the heading angle error (ψ_err) is greater than the first threshold (C ψ,1 ), is chosen to be substantially equal to a heading angle saturation rate ({dot over (ψ)}_lim).

5. The method of claim 4 , wherein, when the absolute value of the heading angle error (ψ_err) is less than the first threshold (C_(ψ,1)) and greater than the second threshold (C ϕ,2 ), the target rate of change ({dot over (ψ)} des ) of the heading angle (ϕ) is computed using the following formula: {dot over (ψ)}_des=sign(ψ_err)√(2(({dot over (ψ)}_lim{circumflex over ( )}2)/(2C_(ψ,1)−C_(ψ,2)))( ψ _err|−C_(ψ,2)/2)), where sign(ψ err ) is a signum function of the heading angle error (ψ_err).

6. The method of claim 4 , wherein, when the absolute value of the heading angle error (ψ_err) is less than the third threshold (C_(ψ,3)), the target rate of change ({dot over (ψ)} des ) of the heading angle (ψ) is chosen to be based on a proportional-integral-derivative control function of the heading angle error (ψ_err).

7. The method of claim 6 , wherein, when the absolute value of the heading angle error (ψ_err) is less than the third threshold (C ψ,3 ), a proportional term in the proportional-integral-derivative control function is substantially equal to (√(({dot over (ψ)}_lim{circumflex over ( )}2)/(C_(ψ,2)(2C_(ψ,1)−C_(ψ,2)))))ψ_err.

8. The method of claim 4 , wherein, when the absolute value of the heading angle error (ψ_err) is less than the second threshold (C_(ψ,2)) and greater than the third threshold (C_(ψ,3)), the target rate of change ({dot over (ψ)} des ) of the heading angle (ψ) is chosen to be substantially proportional to the heading angle error (ψ_err).

9. The method of claim 4 , wherein, when the absolute value of the heading angle error (ψ_err) is less than the second threshold (C_(ψ,2)) and greater than the third threshold (C_(ψ,3)), the target rate of change ({dot over (ψ)} des ) of the heading angle (ψ) is computed according to the following formula: {dot over (ψ)}_des=(√(({dot over (ψ)}_lim{circumflex over ( )}2)/(C_(ψ,2)(2C_(ψ,1)−C_(ψ,2)))))ψ_err.

10. The method of claim 1 , wherein the sliding mode control technique includes using a sigmoid function as a mapping between the target rate of change ({dot over (ψ)} des ) of the heading angle (ψ) and the heading angle error (ψ err ).

11. The method of claim 1 , wherein:
the FL control technique is a first FL control technique;
the sliding mode control technique is a first sliding mode control technique; and
using the commanded bank angle (ϕ cmd ) to control the one or more actuators of the aircraft during flight comprises:
computing a bank angle error (ϕ err ) indicative of a difference between the bank angle (ϕ) of the aircraft and the commanded bank angle (ϕ cmd );
computing a target rate of change ({dot over (ϕ)} des ) for the bank angle (ϕ) of the aircraft using a second sliding mode control technique, an input to the second sliding mode control technique including the bank angle error (ϕ err );
computing a target body roll rate (P c ) for the aircraft using a second FL control technique, an input to the second FL control technique including the target rate of change ({dot over (ϕ)} des ) for the bank angle (ϕ) of the aircraft; and
using the target body roll rate (P c ) to control the one or more actuators of the aircraft during flight.

12. A system for controlling a heading angle (ψ) of an aircraft during flight, the system comprising:
one or more computers operatively coupled to receive one or more signals indicative of a commanded heading angle (ψ cmd ) for the aircraft, the one or more computers being configured to:
compute a heading angle error (ψ_err) indicative of a difference between the heading angle (ψ) of the aircraft and the commanded heading angle (ψ cmd );
compute a target rate of change ({dot over (ψ)} des ) for the heading angle (ψ) of the aircraft using a sliding mode control technique, an input of the sliding mode control technique including the heading angle error (ψ_err);
compute a commanded bank angle (ϕ cmd ) for the aircraft using a feedback linearization (FL) control technique, an input to the FL control technique including the target rate of change ({dot over (ψ)} des ) for the heading angle (ψ) of the aircraft; and
use the commanded bank angle (ϕ cmd ) to control one or more actuators of the aircraft during flight.

13. The system of claim 12 , wherein the FL control technique includes using an inversion of a relationship between the heading angle (ψ) of the aircraft and a bank angle (ϕ) of the aircraft to calculate the commanded bank angle (ϕ cmd ).

14. The system of claim 12 , wherein the FL control technique includes computing the commanded bank angle (ϕ cmd ) using the following formula:

ϕ

Assignees

• Bombardier Inc

Inventors

• Kron Aymeric

Classifications

• G05D1/48

  Control of altitude or depth · CPC title

• G05D1/46

  Control of position or course in three dimensions [3D] · CPC title

• G05D1/49

  Control of attitude, i.e. control of roll, pitch or yaw · CPC title

• G06F17/11

  for solving equations {, e.g. nonlinear equations, general mathematical optimization problems (optimization specially adapted for a specific administrative, business or logistic context G06Q10/04)} · CPC title

• G05D1/085

  to ensure coordination between different movements · CPC title