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

Patent metadata

Field	Value
Publication number	US-12019456-B2
Application number	US-202017038441-A
Country	US
Kind code	B2
Filing date	Sep 30, 2020
Priority date	Sep 30, 2019
Publication date	Jun 25, 2024
Grant date	Jun 25, 2024

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 bank angle (ϕ) of an aircraft during flight, the method comprising:
receiving a commanded bank angle (ϕ cmd ) for the aircraft;
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 sliding mode control technique, an input to the sliding mode control technique including the bank angle error (ϕ err );
computing a target body roll rate (P c ) 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 (ϕ)} des ) for the bank angle (ϕ) of the aircraft; and
using the target body roll rate (P c ) 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 bank angle (ϕ) of the aircraft and one or more body angular rates of the aircraft to calculate the target body roll rate (P c ) for the aircraft.

3. The method of claim 1 , wherein the FL control technique includes computing the target body roll rate (P c ) using the following formula: P c ={dot over (ϕ)} des −tan{circumflex over (θ)} sin{circumflex over (ϕ)}{circumflex over (Q)}−tan {circumflex over (θ)} tan {circumflex over (ϕ)}{circumflex over (R)}, where {circumflex over (θ)} denotes a value indicative of a pitch angle of the aircraft, {circumflex over (ϕ)} denotes a value indicative of the bank angle (ϕ) of the aircraft, {circumflex over (Q)} denotes a value indicative of a body pitch rate of the aircraft, and {circumflex over (R)} denotes a value indicative of a body yaw rate of the aircraft.

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

5. The method of claim 4 , wherein, when the absolute value of the bank angle error (ϕ err ) is less than the first threshold (C ϕ,1 ) and greater than the second threshold (C ϕ,2 ), the target rate of change ({dot over (ϕ)} des ) of the bank angle (ϕ) is computed using the following formula:

ϕ
.

des

=

sign
⁡

(

ϕ

e
⁢
r
⁢
r

)

⁢

2
⁢

k
ϕ
2

⁢

C

ϕ
,
2

⁡

(



ϕ

e
⁢
r
⁢
r



-

c

ϕ
,
2

2

)

,

where sign (ϕ err ) is a signum function of the bank angle error (ϕ err ) and k ϕ denotes a parameter.

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

7. The method of claim 4 , wherein, when the absolute value of the bank angle error (ϕ err ) is less than the second threshold (C ϕ,2 ) and greater than the third threshold (C ϕ,3 ), the target rate of change ({dot over (ϕ)} des ) of the bank angle (ϕ) is chosen to be proportional to the bank angle error (ϕe rr ).

8. The method of claim 4 , wherein, when the absolute value of the bank angle error (ϕ err ) is less than the second threshold (C ϕ,2 ) and greater than the third threshold (C ϕ,3 ), the target rate of change ({dot over (ϕ)} des ) of the bank angle (ϕ) is computed using the following formula: {dot over (ϕ)} des =k ϕ ϕ err , where k ϕ denotes a parameter.

9. 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 (ϕ)} des ) of the bank angle (ϕ) and the bank angle error (ϕ err ).

10. The method as defined in claim 1 , 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 therewith, the program code readable/executable by a computer, processor or logic circuit to perform a method as defined in claim 1 .

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

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