CN107562064B - Attitude control distribution method of aircraft based on multiple actuating mechanisms - Google Patents

Attitude control distribution method of aircraft based on multiple actuating mechanisms Download PDF

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CN107562064B
CN107562064B CN201610512916.4A CN201610512916A CN107562064B CN 107562064 B CN107562064 B CN 107562064B CN 201610512916 A CN201610512916 A CN 201610512916A CN 107562064 B CN107562064 B CN 107562064B
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aircraft
attitude
moment
axis
control
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CN107562064A (en
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秦雷
吴限德
李君龙
谢晓瑛
张锐
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Beijing Institute of Electronic System Engineering
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Beijing Institute of Electronic System Engineering
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Abstract

The invention discloses a multi-actuating mechanism-based attitude control distribution method for an aircraft, which comprises the following steps: s1, establishing an attitude kinematics equation of the aircraft; s2, determining the attitude angle of the aircraft through the attitude transformation matrix; s3, establishing an attitude dynamics equation of the aircraft; s4, calculating the expected control torque of the aircraft according to the control law; and S5, obtaining the state combination of each actuator of the aircraft at the current time by using a control distribution algorithm. The invention is suitable for the maneuvering task requirements of rapid maneuvering and high-precision stability of the large-angle attitude of the aircraft.

Description

Attitude control distribution method of aircraft based on multiple actuating mechanisms
Technical Field
The invention relates to the field of intelligent control of aircrafts. And more particularly, to a multi-actuator based attitude control assignment method for an aircraft.
Background
The aircraft based on multiple actuating mechanisms is a multipurpose and multifunctional novel sheet material combined aircraft which exerts respective advantages and is combined together on the basis of intelligent materials, an intelligent self-adaptive nonlinear control distribution technology and a micro-propulsion system. The aircraft has the advantages of being skin-changeable, capable of greatly improving the performance of orbital transfer maneuver, achieving the best aerodynamic performance, improving the reusability, greatly improving the reconfigurable ability and the like.
The quick maneuver strives for more time for attacking an enemy target and avoiding enemy attack, and the quick high-precision stability after the maneuver is used for ensuring flight stability so as to improve reusability of the aircraft in high-precision flight. Therefore, the rapid maneuvering and high-precision stabilization of the large-angle attitude become one of the key technologies of the aircraft based on the multiple actuating mechanisms.
The choice of actuator determines the maneuverability of the aircraft control system, and the choice of different actuators also determines the maneuverability of the system. Aiming at the complex tasks, a single executing mechanism cannot meet the requirements, the advantages of all mechanism combinations can be exerted in a combined control mode of a plurality of executing mechanisms, the advantage complementation is realized, and the control target is realized at the minimum cost.
Therefore, it is desirable to provide a method for assigning attitude control for a multi-actuator based aircraft.
Disclosure of Invention
The invention aims to provide an attitude control distribution method of an aircraft based on multiple execution mechanisms, which solves the problem of ideal assumed conditions of the traditional attitude control distribution method.
In order to achieve the purpose, the invention adopts the following technical scheme:
a multi-actuator based attitude control distribution method for an aircraft, the method comprising the steps of:
s1, establishing an attitude kinematics equation of the aircraft;
s2, determining the attitude angle of the aircraft through the attitude transformation matrix;
s3, establishing an attitude dynamics equation of the aircraft;
s4, calculating the expected control torque of the aircraft according to the control law;
and S5, obtaining the state combination of each actuator of the aircraft at the current time by using a control distribution algorithm.
Preferably, the attitude kinematics equation of the aircraft in step S1 is:
Figure BDA0001037326360000021
wherein q is0、q1、q2And q is3Quaternion of the attitude of the aircraft at the current moment;
Figure BDA0001037326360000022
and
Figure BDA0001037326360000023
quaternion derivative of aircraft attitude at current time；wxRotating the aircraft around the X axis by the angular velocity; w is ayRotating the aircraft around the Y axis by the angular velocity; w is azThe angular velocity of the aircraft about the Z axis.
Preferably, the posture conversion matrix in step S2 is:
Figure BDA0001037326360000024
preferably, the attitude dynamics equation of the aircraft in step S3 is:
Figure BDA0001037326360000025
wherein, IxRotating inertia around an X axis for the aircraft; i isyRotating inertia around the Y axis for the aircraft; i iszRotating inertia around the Z axis for the aircraft; dwxIs the derivative of the angular speed of rotation of the aircraft about the X axis; dwyIs the derivative of the angular speed of rotation of the aircraft about the Y axis; dwzIs the derivative of the angular speed of rotation of the aircraft about the Z axis; t isxIs the component of all external forces acting on the aircraft on the X axis to the moment of the center of mass; t isyIs the component of the moment of the mass center on the Y axis by all external forces acting on the aircraft; t iszIs the component of the moment of the center of mass to the Z-axis of all external forces acting on the aircraft.
Preferably, step S4 further includes the following sub-steps:
s4.1, calculating quaternion differential of the attitude of the aircraft at the current moment:
dq0＝qc0×q0+qc1×q1+qc2×q2+qc3×q3
dq1＝-qc1×q0+qc1×q1+qc2×q2+qc3×q3
dq2＝-qc3×q1-qc4×q2+qc2×q2+qc3×q3
dq3＝-qc3×q0+qc2×q1-qc1×q2+qc0×q3
wherein dq is0、dq1、dq2And dq3Quaternion differential of the aircraft attitude at the current moment; q. q.sc0、qc1、qc2And q isc3Quaternion for the desired attitude of the aircraft;
s4.2, calculating a nonlinear compensation term T of the expected control moment of the aircraftR：
TR＝J×dwc+w×J×w
Wherein w is the current attitude angular velocity of the aircraft; dwcDesired angular acceleration for the aircraft; j is a matrix of the moment of inertia of the aircraft,
Figure BDA0001037326360000031
s4.3, calculating a proportional term T of the expected control moment of the aircraftp：
Figure BDA0001037326360000032
Wherein, KpIs a proportionality coefficient matrix;
Figure BDA0001037326360000033
s4.4, calculating a derivative term T of the expected control moment of the aircraftd：
Td＝-Kd×J×(w-wc)
Wherein, wcDesired angular velocity for the aircraft;
s4.5, calculating an error compensation term T of the expected control moment of the aircrafte：
Figure BDA0001037326360000034
Wherein ww is the error compensation term normalization coefficient,
Figure BDA0001037326360000035
Kepcompensating the comparison term coefficient for the first error, KedCompensating the comparison term coefficients for the second error;
s4.6, calculating expected control moment T of the aircraftc：
Tc＝TR+Tp+Te+Td。
Preferably, step S5 further includes the following sub-steps:
s5.1, determining a quadrant where the control command is located;
s5.2, selecting an actuating mechanism combination according to the direction of the control instruction;
s5.3, determining possible states of all actuating mechanisms in the actuating mechanism combination;
s5.4, calculating the torque corresponding to the possible state of each actuating mechanism in the actuating mechanism combination;
and S5.5, comparing the state combination of each actuator corresponding to the minimum error value between the sum of the torques corresponding to the possible states of each actuator in the actuator combination and the expected control torque, and taking the state combination of each actuator of the aircraft at the current moment as the state combination of each actuator.
The invention has the following beneficial effects:
according to the technical scheme, under the condition that the execution mechanism has redundancy, optimization distribution under a certain optimization criterion or constraint can be realized, the fault tolerance capability of the system on the fault of the execution mechanism is improved, and a feasible solution is provided for the problems of quick attitude maneuver and high-precision stable control under the condition of redundant heterogeneous execution mechanisms. Specifically, the technical scheme of the invention provides an attitude control method based on MEMS micro-actuators on the basis of establishing an actuator control scheme, establishes an attitude kinematics and a dynamics equation, gives the state of each micro-actuator by formulating a control distribution scheme, namely, determines a dynamic distribution control strategy, sets three typical simulation scenes in simulation, and can be seen from a simulation result that the errors between the current attitude angle and attitude angular velocity and the expected attitude angle and attitude angular velocity are within 0.5 degree, thereby meeting the error requirement, and the startup and shutdown times meet the requirement of the aircraft execution task, thereby realizing the maximum maneuvering angular velocity and the optimal maneuvering task under the constraint of a target function.
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The following describes embodiments of the present invention in further detail with reference to the accompanying drawings.
FIG. 1 illustrates a flow chart of a method for attitude control distribution for a multi-actuator based aircraft.
FIG. 2 illustrates a current attitude angle and desired control angle diagram for an aircraft.
FIG. 3 illustrates a current attitude angular velocity and desired control attitude angular velocity map of an aircraft.
FIG. 4 shows a plot of attitude tracking error for the aircraft after 3 s.
Fig. 5 shows a desired control moment diagram of the aircraft.
Fig. 6 shows an actual output torque diagram of the aircraft.
Fig. 7 shows a schematic diagram of the desired control moment of the aircraft.
Fig. 8 shows a micro-actuator switching diagram for an aircraft.
Detailed Description
In order to more clearly illustrate the invention, the invention is further described below with reference to preferred embodiments and the accompanying drawings. Similar parts in the figures are denoted by the same reference numerals. It is to be understood by persons skilled in the art that the following detailed description is illustrative and not restrictive, and is not to be taken as limiting the scope of the invention.
As shown in fig. 1, the method for allocating attitude control to an aircraft based on multiple actuators (hereinafter referred to as an aircraft) according to this embodiment includes the following steps:
s1, establishing an attitude kinematics equation of the aircraft;
s2, determining the attitude angle of the aircraft through the attitude transformation matrix;
s3, establishing an attitude dynamics equation of the aircraft;
s4, calculating the expected control torque of the aircraft according to the control law;
and S5, obtaining the state combination of each actuator of the aircraft at the current time by using a control distribution algorithm.
Wherein the content of the first and second substances,
the attitude kinematics equation of the aircraft in step S1 is:
Figure BDA0001037326360000051
wherein q is0、q1、q2And q is3The quaternion of the attitude (from an inertial coordinate system to a projectile coordinate system) of the aircraft at the current moment;
Figure BDA0001037326360000052
and
Figure BDA0001037326360000053
the quaternion derivative of the aircraft attitude at the current moment; w is axThe angular velocity of rotation of the aircraft about the X axis (inertial frame); w is ayRotating the aircraft around the Y axis by the angular velocity; w is azThe angular velocity of the aircraft about the Z axis.
The specific process of step S2 is:
and rotating the inertial coordinate system according to the Z-Y-X sequence to convert the attitude quaternion of the aircraft into an Euler angle of the aircraft. Attitude transformation matrix
Figure BDA0001037326360000054
Comprises the following steps:
Figure BDA0001037326360000055
the attitude dynamics equation of the aircraft in step S3 is:
Figure BDA0001037326360000056
wherein, IxAs aircraftRotational inertia about the X axis; i isyRotating inertia around the Y axis for the aircraft; i iszRotating inertia around the Z axis for the aircraft; dwxIs the derivative of the angular speed of rotation of the aircraft about the X axis; dwyIs the derivative of the angular speed of rotation of the aircraft about the Y axis; dwzIs the derivative of the angular speed of rotation of the aircraft about the Z axis; t isxThe component of all external forces (including aerodynamic force and thrust) acting on the aircraft to the moment of the mass center on the X axis (a missile coordinate system); t isyIs the component of all external forces (including aerodynamic force and thrust) acting on the aircraft to the moment of the center of mass on the Y axis; t iszIs the component of the moment of mass in the Z-axis for all external forces (including aerodynamic forces and thrust) acting on the aircraft.
Step S4 "calculate the desired control moment of the aircraft according to the control law" further comprises the sub-steps of:
s4.1, calculating quaternion differential of the attitude of the aircraft at the current moment:
dq0＝qc0×q0+qc1×q1+qc2×q2+qc3×q3
dq1＝-qc1×q0+qc1×q1+qc2×q2+qc3×q3
dq2＝-qc3×q1-qc4×q2+qc2×q2+qc3×q3
dq3＝-qc3×q0+qc2×q1-qc1×q2+qc0×q3
wherein dq is0、dq1、dq2And dq3Quaternion differential of the aircraft attitude at the current moment; q. q.sc0、qc1、qc2And q isc3Quaternion for the desired attitude of the aircraft;
s4.2, calculating a nonlinear compensation term T of the expected control moment of the aircraftR：
TR＝J×dwc+w×J×w
Wherein w is the current attitude angular velocity of the aircraft; dwcDesired angular acceleration for the aircraft; j is a matrix of the moment of inertia of the aircraft,
Figure BDA0001037326360000061
s4.3, calculating a proportional term T of the expected control moment of the aircraftp：
Figure BDA0001037326360000062
Wherein, KpIs a proportionality coefficient matrix;
Figure BDA0001037326360000063
s4.4, calculating a derivative term T of the expected control moment of the aircraftd：
Td＝-Kd×J×(w-wc)
Wherein, wcDesired angular velocity for the aircraft;
s4.5, calculating an error compensation term T of the expected control moment of the aircrafte：
Figure BDA0001037326360000064
Wherein ww is the error compensation term normalization coefficient,
Figure BDA0001037326360000065
Kepcompensating the comparison term coefficient for the first error, KedCompensating the comparison term coefficients for the second error;
s4.6, calculating expected control moment T of the aircraftc：
Tc＝TR+Tp+Te+Td。
The specific process of step S5 "obtaining the state combination of each actuator of the aircraft at the current time by using the control distribution algorithm" includes:
knowing the normal vector of the convex cone:
Figure BDA0001037326360000066
the combination of actuators with an extended geometry to determine a 50 degree mounting angle configuration is shown in tables 1-14, and pre-bound data can be further determined based on the distribution of the moments generated by the actuators in three-dimensional space: the normal vector of the conical surface of the multi-cone body formed by the combination of the actuating mechanisms, the actuating mechanism combination contained in each quadrant and the positive and negative conditions of dot products which are required to be met when the expected control moment belongs to the inner part of the area formed by the combination of the actuating mechanisms. The results of the calculations are shown in tables 1-15 and tables 1-16, respectively:
tables 1-14 results of determining actuator combination tables based on extended geometry
Figure BDA0001037326360000071
Table 1-15 cone normal vector of multi-cone formed by combining actuating mechanisms
Conical surface Normal vector Conical surface Normal vector Conical surface Normal vector
<T7,T1> n1＝[0.8609,-0.5068,0.0443]T <T8,T6> n5＝[-0.8609,-0.5068,-0.0443]T <T1,T8> n9＝[0.2042,-0.6922,0.6922]T
<T1,T3> n2＝[0.8609,-0.0443,-0.5068]T <T2,T8> n6＝[-0.8609,0.0443,-0.5068]T <T3,T2> n10＝[0.2042,-0.6922,-0.6922]T
<T3,T5> n3＝[0.8609,0.5068,-0.0443]T <T4,T2> n7＝[-0.8609,0.5068,0.0443]T <T5,T4> n11＝[0.2042,0.6922,-0.6922]T
<T5,T7> n4＝[0.8609,0.0443,0.5068]T <T6,T4> n8＝[-0.8609,-0.0443,0.5068]T <T7,T6> n12＝[0.2042,0.6922,0.6922]T
Tables 1 to 15<Ti,Tj>Representing the cone formed by the moment vectors of the ith and jth actuators, in practice care should be taken that the direction of the normal is consistent with the decision logic in tables 1-16.
Actuator combinations and decision logic contained in quadrants of tables 1-16
Figure BDA0001037326360000072
Figure BDA0001037326360000081
(u) in tables 1 to 163d,ni) Representing the desired control moment u3dNormal vector n of the sum planeiDot product of (1), and symbol "&"represents a logical and. From tables 1-16, it can be seen that for the 50 degree thrust configuration, 12 cone direction vectors, n respectively, need to be stored when pre-setting the data1、n2、n3、n4、n5、n6、n7、n8、n9、n10、n11And n12. Each quadrant may contain three possible actuator combinations, and three dot products may be calculated when determining the optimal actuator combination.
The attitude angular response curve is shown in fig. 2, the attitude angular velocity response curve is shown in fig. 3, the attitude tracking error curve after 3s is shown in fig. 4, the expected control torque curve is shown in fig. 5, the actual output torque curve of the actuator is shown in fig. 6, the expected control torque curve after 5s and the actual output torque curve of the actuator are shown in fig. 7, and the switching state curves of 8 actuators are shown in fig. 8.
The optimal actuating mechanism combination:
Figure BDA0001037326360000091
step S5 "obtaining the status combination of each actuator of the aircraft at the current time using the control distribution algorithm" further includes the sub-steps of:
s5.1, determining a quadrant where the control command is located:
setting quadrant symbol as num, and the identifier flag of the expected control moment of the aircraft, satisfying:
Figure BDA0001037326360000092
num＝num+flag×2i-1(i＝1，2，3)
Figure BDA0001037326360000093
s5.2, selecting an actuating mechanism combination according to the direction of the control command:
setting (u)3d,n1) Is represented by the formula (1), (u)3d,n2) Is represented by the formula (2), (u)3d,n9) Is represented by the formula (3), (u)3d,n3) Is represented by the formula (4), (u)3d,n10) Is represented by the formula (5), (u)3d,n4) Is represented by the formula (6), (u)3d,n11) Is represented by the formula (7), (u)3d,n12) Is represented by the formula (8), (u)3d,n5) Is represented by the formula (9), (u)3d,n6) Is represented by the formula (10), (u)3d,n7) Is represented by the formula (11), (u)3d,n8) Is represented by the formula (12).
If num is 7:
if equation (1) is greater than or equal to 0 and equation (2) is greater than or equal to 0, the actuator combination is [ 1357 ]
If equation (1) is <0 and equation (3) is <0, then the combination of actuators is [ 1678 ]
If the formula (2) is less than 0 and the formula (3) is greater than or equal to 0, the actuator combination is [ 1238 ]
If num is 6:
if the formula (2) is more than or equal to 0 and the formula (4) is more than or equal to 0, the combination of the execution mechanisms is [ 1357 ]
If equation (2) is <0 and equation (5) is <0, then the combination of actuators is [ 1238 ]
If the formula (4) is less than 0 and the formula (5) is more than or equal to 0, the actuator combination is [ 2345 ]
If num is 4:
if the formula (4) is greater than or equal to 0 and the formula (6) is greater than or equal to 0, the actuator combination is [ 1357 ]
If equation (4) is <0 and equation (7) is <0, the combination of actuators is [ 2345 ]
If the formula (6) is less than 0 and the formula (7) is more than or equal to 0, the combination of the execution mechanisms is [ 4567 ]
If num is 5:
if the formula (6) is greater than or equal to 0 and the formula (1) is greater than or equal to 0, the actuator combination is [ 1357 ]
If equation (6) <0 and equation (8) <0, the combination of actuators is [ 4567 ]
If the formula (1) is less than 0 and the formula (8) is greater than or equal to 0, the combination of the execution mechanisms is [ 1678 ]
If num is 3:
if the formula (9) is greater than or equal to 0 and the formula (10) is greater than or equal to 0, the actuator combination is [ 2468 ]
If equation (9) is <0 and equation (3) is <0, then the combination of actuators is [ 1238 ]
If the formula (10) is less than 0 and the formula (3) is more than or equal to 0, the actuator combination is [ 2345 ]
If num is 2:
if equation (10) is greater than or equal to 0 and equation (11) is greater than or equal to 0, the actuator combination is [ 2468 ]
If equation (10) is <0 and equation (5) is <0, then the combination of actuators is [ 1238 ]
If equation (11) is less than 0 and equation (5) is greater than or equal to 0, the actuator combination is [ 2345 ]
If num is 0:
if equation (11) is greater than or equal to 0 and equation (12) is greater than or equal to 0, the actuator combination is [ 2468 ]
If equation (11) is <0 and equation (7) is <0, the combination of actuators is [ 2345 ]
If the formula (12) is less than 0 and the formula (7) is more than or equal to 0, the combination of the execution mechanisms is [ 4567 ]
If num is 1:
if equation (12) is greater than or equal to 0 and equation (9) is greater than or equal to 0, the actuator combination is [ 2468 ]
If equation (12) is <0 and equation (8) is <0, the combination of actuators is [ 4567 ]
If equation (9) is <0 and equation (8) is greater than or equal to 0, the combination of execution mechanisms is [ 1678 ].
S5.3, determining possible states of all actuating mechanisms in the actuating mechanism combination
Calculating all possible switch states under 16 conditions (8 thrusters, each thruster has two states of on and off), and setting the control command of the jth actuator as U0(j) The ith switch combination UU (i, j) of the jth actuator converts the 0-15 combinations into a four-digit binary number:
Figure BDA0001037326360000111
abcd is a control command with four bits and binary number (i is 0-15).
S5.4, calculating the torque corresponding to the possible state of each actuating mechanism in the actuating mechanism combination:
if U is0(j) 0 and UU (i, j) 0, then:
F(j)＝0
if U is0(j) 1 and UU (i, j) ═ 1, then:
F(j)＝fmax
wherein f ismaxThe maximum thrust of the actuating mechanism;
if U is0(j) Not equal to 1 or UU (i, j) ≠ 1
F(j)＝0.5×fmax
T(:,i)＝M×F'(j)
In the above formula, M is an actuator mounting matrix (representing a control torque matrix, and the control torques in the pitch, yaw and roll directions are obtained by multiplying the forces borne by eight actuators by the torque, which can indicate 8 actuators), f (j) is the torque generated by the jth actuator, and T (: i) is the sum of the control torques generated in 16 switching states.
S5.5, comparing to obtain the state combination of each actuating mechanism corresponding to the minimum error value between the sum of the torques corresponding to the possible states of each actuating mechanism in the actuating mechanism combination and the expected control torque, and taking the state combination of each actuating mechanism of the aircraft at the current moment as the state combination of each actuating mechanism:
setting error | | | T (: 1) -Tc||，
If | | | T (: i) -Tc||＜error
Then error | | | T (: i) -Tc||(i＝0-15)
And (4) finding errors between the sum of the moments generated by each actuating mechanism and the expected control moment through traversal, and recording the combination serial number in the ith switch state corresponding to the minimum error value.
The invention can realize high-precision attitude control based on multiple actuating mechanisms, and has the outstanding characteristics of exerting the combined control advantage of the multiple actuating mechanisms, thereby realizing high-precision attitude control and meeting the requirement of rapidly switching an aircraft between control modes such as attitude maneuver, high-precision stability and the like.
It should be understood that the above-mentioned embodiments of the present invention are only examples for clearly illustrating the present invention, and are not intended to limit the embodiments of the present invention, and it will be obvious to those skilled in the art that other variations or modifications may be made on the basis of the above description, and all embodiments may not be exhaustive, and all obvious variations or modifications may be included within the scope of the present invention.

Claims (2)

1. A method for distributing attitude control of an aircraft based on multiple actuators is characterized by comprising the following steps:
s1, establishing an attitude kinematics equation of the aircraft;
s2, determining the attitude angle of the aircraft through the attitude transformation matrix;
s3, establishing an attitude dynamics equation of the aircraft;
s4, calculating the expected control torque of the aircraft according to the control law;
s5, obtaining the state combination of each actuating mechanism of the aircraft at the current time by using a control distribution algorithm;
wherein, the attitude kinematics equation of the aircraft in the step S1 is:
Figure FDA0002415909090000011
wherein q is0、q1、q2And q is3Quaternion of the attitude of the aircraft at the current moment;
Figure FDA0002415909090000012
and
Figure FDA0002415909090000013
the quaternion derivative of the aircraft attitude at the current moment; w is axRotating the aircraft around the X axis by the angular velocity; w is ayRotating the aircraft around the Y axis by the angular velocity; w is azRotating the aircraft around the Z axis by an angular velocity;
the attitude dynamics equation of the aircraft in step S3 is:
Figure FDA0002415909090000014
wherein, IxRotating inertia around an X axis for the aircraft; i isyRotating inertia around the Y axis for the aircraft; i iszRotating inertia around the Z axis for the aircraft; dwxIs the derivative of the angular speed of rotation of the aircraft about the X axis; dwyIs the derivative of the angular speed of rotation of the aircraft about the Y axis; dwzIs the derivative of the angular speed of rotation of the aircraft about the Z axis; t isxIs the component of all external forces acting on the aircraft on the X axis to the moment of the center of mass; t isyIs the component of the moment of the mass center on the Y axis by all external forces acting on the aircraft; t iszIs the component of all external forces acting on the aircraft to the moment of the center of mass on the Z axis;
step S4 further includes the following sub-steps:
s4.1, calculating quaternion differential of the attitude of the aircraft at the current moment:
dq0＝qc0×q0+qc1×q1+qc2×q2+qc3×q3
dq1＝-qc1×q0+qc1×q1+qc2×q2+qc3×q3
dq2＝-qc3×q1-qc4×q2+qc2×q2+qc3×q3
dq3＝-qc3×q0+qc2×q1-qc1×q2+qc0×q3
wherein dq is0、dq1、dq2And dq3Quaternion differential of the aircraft attitude at the current moment; q. q.sc0、qc1、qc2And q isc3Quaternion for the desired attitude of the aircraft;
s4.2, calculating a nonlinear compensation term T of the expected control moment of the aircraftR：
TR＝J×dwc+w×J×w
Wherein w is the current attitude angular velocity of the aircraft; dwcDesired angular acceleration for the aircraft; j is a matrix of the moment of inertia of the aircraft,
Figure FDA0002415909090000021
s4.3, calculating a proportional term T of the expected control moment of the aircraftp：
Figure FDA0002415909090000022
Wherein, KpIs a proportionality coefficient matrix;
Figure FDA0002415909090000023
s4.4, calculating a derivative term T of the expected control moment of the aircraftd：
Td＝-Kd×J×(w-wc)
Wherein, wcDesired angular velocity for the aircraft;
s4.5, calculating an error compensation term T of the expected control moment of the aircrafte：
Figure FDA0002415909090000024
Wherein ww is the error compensation term normalization coefficient,
Figure FDA0002415909090000025
Kepcompensating the comparison term coefficient for the first error, KedCompensating the comparison term coefficients for the second error;
s4.6, calculating expected control moment T of the aircraftc：
Tc＝TR+Tp+Te+Td；
Step S5 further includes the following sub-steps:
s5.1, determining a quadrant where the control command is located;
s5.2, selecting an actuating mechanism combination according to the direction of the control instruction;
s5.3, determining possible states of all actuating mechanisms in the actuating mechanism combination;
s5.4, calculating the torque corresponding to the possible state of each actuating mechanism in the actuating mechanism combination;
and S5.5, comparing the state combination of each actuator corresponding to the minimum error value between the sum of the torques corresponding to the possible states of each actuator in the actuator combination and the expected control torque, and taking the state combination of each actuator of the aircraft at the current moment as the state combination of each actuator.
2. The method according to claim 1, wherein the gesture transformation matrix in step S2 is:
Figure FDA0002415909090000031
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