CN112486218A - Helicopter control method and system - Google Patents

Helicopter control method and system Download PDF

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Publication number
CN112486218A
CN112486218A CN202011387658.4A CN202011387658A CN112486218A CN 112486218 A CN112486218 A CN 112486218A CN 202011387658 A CN202011387658 A CN 202011387658A CN 112486218 A CN112486218 A CN 112486218A
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helicopter
calculated
pitch
partial derivative
control quantity
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CN112486218B (en
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田刚印
邓海波
陈佳
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Shenzhen Lianhe Airplane Technology Co ltd
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Shenzhen Lianhe Airplane Technology Co ltd
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    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D13/00—Control of linear speed; Control of angular speed; Control of acceleration or deceleration, e.g. of a prime mover
    • G05D13/62—Control of linear speed; Control of angular speed; Control of acceleration or deceleration, e.g. of a prime mover characterised by the use of electric means, e.g. use of a tachometric dynamo, use of a transducer converting an electric value into a displacement

Abstract

The application discloses a control method and a system of a helicopter, wherein the control method comprises the following steps: the following steps are executed in a circulating way: calculating a feedforward control quantity in real time according to the off-line pneumatic data; calculating a controlled variable error between the feedforward controlled variable and the real-time controlled variable according to the real-time controlled variable; calculating feedback control quantity according to the real-time navigation measurement data and the flight reference quantity; and determining an output control quantity according to the feedforward control quantity, the control quantity error and the feedback control quantity, and taking the output control quantity as a rudder instruction of the helicopter. The method adopts a real-time estimation compensation control method, and combines the feedforward control quantity and the feedback control quantity obtained by real-time navigation measurement data to improve the wind resistance control performance of the helicopter, so that the helicopter is ensured to always work near the actual trim state.

Description

Helicopter control method and system
Technical Field
The application relates to the technical field of helicopter control, in particular to a helicopter control method and system.
Background
The existing helicopter flight control wind resistance control method mainly adopts Proportional-Integral-Derivative (PID) control, and improves the wind resistance of the helicopter by means of Integral effect and large gain, which puts high requirements on single-machine equipment of the system, such as steering engine response speed, inertial measurement unit measurement accuracy and the like. Therefore, a helicopter wind-resistant control method which reduces the requirements of single-machine equipment and is low in control difficulty is urgently needed to be found.
Disclosure of Invention
The application aims to provide a control method and a control system of a helicopter, which can reduce the channel coupling of the helicopter and improve the response speed of flight control of the helicopter.
The application provides a control method of a helicopter, comprising the following steps: the following steps are executed in a circulating way: calculating a feedforward control quantity according to the off-line pneumatic data; calculating a controlled variable error between the feedforward controlled variable and the real-time controlled variable according to the real-time controlled variable; calculating feedback control quantity according to the real-time measurement data; and determining an output control quantity according to the feedforward control quantity, the control quantity error and the feedback control quantity, and taking the output control quantity as a rudder instruction of the helicopter.
Preferably, the output control amount is an integrated sum of the feedforward control amount, the control amount error, and the feedback control amount.
Preferably, the feedback control amount is calculated using the following formula:
Figure BDA0002811496090000021
wherein, Vrefy(k) The reference side flying speed, V, of the helicopter at the current moment ky(k) K actual side flight speed of helicopter at current moment, ay(k) The actual lateral flight acceleration of the helicopter at the current moment k, and gamma (k) is the roll attitude angle, omega, of the helicopter at the current moment kxb(k) At the current moment k, the roll angular velocity, V, of the helicopterrefx(k) Reference forward flight speed, V, of helicopter at current time kx(k) K actual forward flight speed of helicopter at current moment, ax(k) The actual forward flight acceleration of the helicopter at the current moment k, and theta (k) is the pitch attitude angle of the helicopter at the current moment k, omegayb(k) Is the pitch angle velocity, V, of the helicopter at the current moment krefz(k) Is the reference climbing speed, h, of the helicopter at the current moment kref(k) Is the reference flight height of k helicopter at the current moment, h (k) is the actual flight height of k helicopter at the current moment, Vz(k) K actual climbing speed of helicopter at present moment, az(k) Is the actual climbing acceleration, omega, of the helicopter at the current moment krefzb(k) Is the reference course angular velocity, psi, of the helicopter at the current time kref(k) Is the reference course angle, omega, of the helicopter at the current moment kzb(k) Is the actual course angular velocity of k helicopter at the current moment, psi (k) is the actual course angle of k helicopter at the current moment, u0f(k)、u1f(k)、u2f(k)、u3f(k) Respectively are feedback control quantities of a roll channel, a pitch channel, a high-direction channel and a course channel of the k helicopter at the current moment.
Preferably, the feedforward control quantity comprises at least the lateral cyclic pitch, the longitudinal cyclic pitch, the collective pitch and the tail pitch of the helicopter.
Preferably, the feedforward control amount is calculated according to solving a first function
Figure BDA0002811496090000022
Wherein V is the offline flight speed of the helicopter, theta is the offline pitch attitude angle of the helicopter, gamma is the offline roll attitude angle of the helicopter, and FZbFor total off-line resultant force acting vertically on the helicopter body, MXb、MYb、MZbRespectively, the total moment of the offline, u, acting on the helicopter body in the direction X, Y, Z0For transverse cyclic pitch, u, of helicopters1For longitudinal cyclic pitch, u, of helicopters2Is the collective pitch u of the helicopter3Is the tail pitch of the helicopter.
Preferably, the control quantity error includes a lateral cyclic pitch error Δ u0(k) Longitudinal cyclic pitch error Δ u1(k) Total distance error Δ u2(k) Tail rotor pitch error delta u3(k)；
The following formula is adopted to calculate the transverse periodic variable pitch error delta u0(k)
e0＝u0(k)-z01(k-1)
Wherein z is01(k-1)＝u0_o(k-1)
z01(k)＝z01(k-1)+[β01e0+z02(k-1)]T
z02(k)＝z02(k-1)+[A0z02(k-1)-B0z01(k-1)+C0u0(k-1)+z03(k-1)+β02]T
z03(k)＝z03(k-1)+β03e0
Δu0(k)＝z03(k)/D0
Wherein T is a control period, z01(k) Is an estimated value of the roll angular velocity of the k helicopter at the current moment, z02(k) Is an estimated value of roll angular acceleration of the helicopter at the current moment k, z03(k) Is the first interference estimate, z, of k helicopters at the current time01(k-1) is the roll angular velocity of the helicopter at the previous moment k-1, z02(k-1) roll angular acceleration of the helicopter, z, at the previous moment k-103(k-1) is the first interference value, beta, of the helicopter at the last moment k-101、β02、β03To a specified value, A0、B0、C0Respectively, a first characteristic quantity identification parameter, D, of the helicopter0Identifying a parameter, u, for a first model of a helicopter0(k) For the lateral cyclic variation of k helicopter at the present moment, u0_o(k-1) is the output lateral cyclic variation of the helicopter at the previous moment k-1, e0An error is estimated for the roll control amount.
Preferably, the longitudinal cyclic pitch error Δ u is calculated using the following formula1(k)
e1＝u1(k)-z11(k-1)
Wherein z is11(k-1)＝u1_o(k-1)
z11(k)＝z11(k-1)+[β11e0+z12(k-1)]T
z12(k)＝z12(k-1)+[A1z12(k-1)-B1z11(k-1)+C1u1(k-1)+z13(k-1)+β12]T
z13(k)＝z13(k-1)+β13e1
Δu1(k)＝z13(k)/D1
Wherein T is a control period, z11(k) Is an estimated value of the pitch angle speed of the k helicopter at the current moment, z12(k) Is an estimated value of the pitch angle acceleration of the k helicopter at the current moment, z13(k) Is a second interference estimate, z, for the current time k helicopter11(k-1) is the pitch angular velocity of the helicopter at the previous moment k-1, z12(k-1) is the pitch acceleration of the helicopter at the previous moment k-1, z13(k-1) is the second interference value, beta, of the helicopter at the last moment k-111、β12、β13To a specified value, A1、B1、C1Respectively, a second characteristic quantity identification parameter, D, of the helicopter1Identifying a parameter, u, for a second model of the helicopter1(k) For the lateral cyclic variation of k helicopter at the present moment, u1_o(k-1) is the output lateral cyclic variation of the helicopter at the previous moment k-1, e1An error is estimated for the pitch control amount.
Preferably, the first function is solved by using a gradient method for the following second function
Figure BDA0002811496090000041
Wherein f isi 2(x) Representing the square of the target error function term.
Preferably, the first function is solved using the following formula
x(k+1)＝x(k)-[f′(x(k))]-1f(x(k))
Figure BDA0002811496090000042
Wherein x is(k+1)Representing the value of the parameter to be found next, x(k)Representing the current value of the parameter to be found, f' (x)(k)) Jacobian matrix, f (x), representing the error function versus the unknown quantity(k)) Representing a solution of a first functionCalculating the value of the current,
Figure BDA0002811496090000043
representing the total off-line resultant force F acting on the vertical direction of the helicopter body in the first functionZbFor transverse cyclic variation u of helicopter0The partial derivative is calculated and the partial derivative is calculated,
Figure BDA0002811496090000044
representing the total off-line resultant force F acting on the vertical direction of the helicopter body in the first functionZbFor longitudinal cyclic variation u of helicopter1The partial derivative is calculated and the partial derivative is calculated,
Figure BDA0002811496090000045
representing the total off-line resultant force F acting on the vertical direction of the helicopter body in the first functionZbTotal distance u to helicopter2The partial derivative is calculated and the partial derivative is calculated,
Figure BDA0002811496090000046
representing the total off-line resultant force F acting on the vertical direction of the helicopter body in the first functionZbTail rotor pitch u for helicopter3Calculating a partial derivative;
Figure BDA0002811496090000051
representing the sum of the moments M taken off-line in the first function and acting in the X direction on the helicopter bodyXbFor transverse cyclic variation u of helicopter0The partial derivative is calculated and the partial derivative is calculated,
Figure BDA0002811496090000052
representing the sum of the moments M taken off-line in the first function and acting in the X direction on the helicopter bodyXbFor longitudinal cyclic variation u of helicopter1The partial derivative is calculated and the partial derivative is calculated,
Figure BDA0002811496090000053
representing the sum of the moments M taken off-line in the first function and acting in the X direction on the helicopter bodyXbTotal distance u to helicopter2The partial derivative is calculated and the partial derivative is calculated,
Figure BDA0002811496090000054
representing the sum of the moments M taken off-line in the first function and acting in the X direction on the helicopter bodyXbTail rotor pitch u for helicopter3Calculating a partial derivative;
Figure BDA0002811496090000055
representing the sum of the moments M taken off-line in the first function, acting in the direction Y on the helicopter bodyYbFor transverse cyclic variation u of helicopter0The partial derivative is calculated and the partial derivative is calculated,
Figure BDA0002811496090000056
representing the sum of the moments M taken off-line in the first function, acting in the direction Y on the helicopter bodyYbLongitudinal cyclic variable pitch u of medium-to-vertical lift1The partial derivative is calculated and the partial derivative is calculated,
Figure BDA0002811496090000057
representing the sum of the moments M taken off-line in the first function, acting in the direction Y on the helicopter bodyXbTotal distance u to helicopter2The partial derivative is calculated and the partial derivative is calculated,
Figure BDA0002811496090000058
representing the sum of the moments M taken off-line in the first function, acting in the direction Y on the helicopter bodyXbTail rotor pitch u for helicopter3Calculating a partial derivative;
Figure BDA0002811496090000059
representing the total moment M of the first function in the Z direction acting on the helicopter bodyZbFor transverse cyclic variation u of helicopter0The partial derivative is calculated and the partial derivative is calculated,
Figure BDA00028114960900000510
representing the total moment M of the first function in the Z direction acting on the helicopter bodyZbLongitudinal cyclic variable pitch u of medium-to-vertical lift1The partial derivative is calculated and the partial derivative is calculated,
Figure BDA00028114960900000511
representing a first functionOff-line total moment M acting on helicopter body in Z directionZbTotal distance u to helicopter2The partial derivative is calculated and the partial derivative is calculated,
Figure BDA00028114960900000512
representing the total moment M of the first function in the Z direction acting on the helicopter bodyZbTail rotor pitch u for helicopter3And (5) calculating a partial derivative.
The application also provides a control system of the helicopter, which comprises a feedforward control quantity calculation module, a control quantity error calculation module, a feedback control quantity calculation module and a rudder instruction determination module; the feedforward control quantity calculation module calculates the feedforward control quantity in real time according to the off-line pneumatic data; the control quantity error calculation module calculates a control quantity error between the feedforward control quantity and the real-time control quantity according to the real-time control quantity; the feedback control quantity calculation module calculates the feedback control quantity according to the real-time navigation measurement data and the flight reference quantity; and the rudder instruction determining module determines an output control quantity according to the feedforward control quantity, the control quantity error and the feedback control quantity, and takes the output control quantity as a rudder instruction of the helicopter.
Drawings
In order to more clearly illustrate the embodiments of the present application or the technical solutions in the prior art, the drawings needed to be used in the description of the embodiments or the prior art will be briefly described below, it is obvious that the drawings in the following description are only some embodiments described in the present application, and other drawings can be obtained by those skilled in the art according to the drawings.
FIG. 1 is a flow chart of a method of controlling a helicopter provided by the present application;
FIG. 2 is a flow chart for solving for feedforward control quantities as provided herein;
FIG. 3 is a schematic diagram of the calculation of lateral cyclic variation error provided herein;
fig. 4 is a block diagram of a control system of the helicopter provided in the present application.
Detailed Description
The technical solutions in the embodiments of the present application are clearly and completely described below with reference to the drawings in the embodiments of the present application, and it is obvious that the described embodiments are some, but not all, embodiments of the present application. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present application.
Example one
The application provides a control method of a helicopter. Fig. 1 is a flowchart of a control method of a helicopter provided by the present application.
As shown in fig. 1, the helicopter control method comprises the following steps:
s110: according to the characteristics of the helicopter, the existing off-line pneumatic data is utilized to calculate the off-line balancing quantity to be used as the feedforward control quantity. The off-line pneumatic data is obtained through wind tunnel test or CFD simulation.
Specifically, the feedforward control amount is calculated according to solving a first function
Figure BDA0002811496090000061
Wherein V is the offline flight speed of the helicopter, theta is the offline pitch attitude angle of the helicopter, gamma is the offline roll attitude angle of the helicopter, and FZbFor total off-line resultant force acting vertically on the helicopter body, MXb、MYb、MZbRespectively, the total moment of the offline, u, acting on the helicopter body in the direction X, Y, Z0For transverse cyclic pitch, u, of helicopters1For longitudinal cyclic pitch, u, of helicopters2Is the collective pitch u of the helicopter3Is the tail pitch of the helicopter. Wherein the feedforward control quantity comprises u0、u1、u2、u3As unknown. V, theta, gamma, FZb，MXb，MYb，MZbIn known amounts.
Compared with the existing full equation, the first function reduces two equations, the solution parameters are few, the solution method is easy, and a complex nonlinear differential equation set does not need to be solved.
As one embodiment, the first function is solved by an iteration method, and the off-line balancing quantity required to act on the airplane is calculated in a reverse mode.
As shown in fig. 2, solving the first function includes the steps of:
s210: iterating the following second function by gradient method
Figure BDA0002811496090000071
Wherein f isi 2(x) Representing the square of the target error function term.
From the initial point
Figure BDA0002811496090000072
Starting along a second function
Figure BDA0002811496090000073
Is decreased in the direction (i.e., negative gradient direction) in which the value of (c) is decreased most rapidly, gradually
Figure BDA0002811496090000074
The value is obtained.
S220: judging whether the latest solution value is less than the first extreme value epsilon1(ii) a If yes, go to S260; otherwise, step S230 is performed.
S230: judging whether the iteration times of the second function are smaller than the specified times or not; if yes, returning to S210; otherwise, step S240 is performed.
S240: iterating using a third function
x(k+1)＝x(k)-[f′(x(k))]-1f(x(k)) (3)
Figure BDA0002811496090000075
Wherein x is(k+1)Representing the value of the parameter to be found next, x(k)Represents the value of the parameter to be currently sought,f′(x(k)) Jacobian matrix, f (x), representing the error function versus the unknown quantity(k)) A resolved value representing a first function is expressed,
Figure BDA0002811496090000081
and the coefficient of the Jacobian matrix is represented and is the partial derivative value of each equation item in the first function to each parameter to be solved.
In particular, the amount of the solvent to be used,
Figure BDA0002811496090000082
representing F in a first functionZbFor u is paired0The partial derivative is calculated and the partial derivative is calculated,
Figure BDA0002811496090000083
representing F in a first functionZbFor u is paired1The partial derivative is calculated and the partial derivative is calculated,
Figure BDA0002811496090000084
representing F in a first functionZbThe partial derivative is taken over u2,
Figure BDA0002811496090000085
representing F in a first functionZbFor u is paired3Calculating a partial derivative;
Figure BDA0002811496090000086
representing M in a first functionXbFor u is paired0The partial derivative is calculated and the partial derivative is calculated,
Figure BDA0002811496090000087
representing M in a first functionXbFor u is paired1The partial derivative is calculated and the partial derivative is calculated,
Figure BDA0002811496090000088
representing M in a first functionXbThe partial derivative is taken over u2,
Figure BDA0002811496090000089
representing M in a first functionXbFor u is paired3Calculating a partial derivative;
Figure BDA00028114960900000810
representing M in a first functionYbFor u is paired0The partial derivative is calculated and the partial derivative is calculated,
Figure BDA00028114960900000811
representing a first function MYbMiddle pair of u1The partial derivative is calculated and the partial derivative is calculated,
Figure BDA00028114960900000812
representing M in a first functionXbThe partial derivative is taken over u2,
Figure BDA00028114960900000813
representing M in a first functionXbFor u is paired3Calculating a partial derivative;
Figure BDA00028114960900000814
representing M in a first functionZbFor u is paired0The partial derivative is calculated and the partial derivative is calculated,
Figure BDA00028114960900000815
representing a first function MZbMiddle pair of u1The partial derivative is calculated and the partial derivative is calculated,
Figure BDA00028114960900000816
representing M in a first functionZbThe partial derivative is taken over u2,
Figure BDA00028114960900000817
representing u in a first functionZbFor u is paired3And (5) calculating a partial derivative.
The initial value of the third function is the latest solution value of the second function.
S250: judgment max (Δ u)i(k) Whether or not it is less than the second extreme value ε2(ii) a If yes, go to S260; otherwise, return to step S240. Wherein, Δ ui(k) Represents the ith (i belongs to {0,1,2,3}) feedforward control quantity u at the current time ki(k) The error of (a) is detected,
Figure BDA00028114960900000818
represents the maximum value among the errors of the four feedforward control amounts.
S260: the latest solution value is output as feedforward control quantity, and u is output0、u1、u2、u3Feedforward lateral cyclic variation u of helicopter as current moment k0(k) Feed forward longitudinal cyclic variation u1(k) Feedforward total distance u2(k) And feed forward tail pitch u3(k)。
As another example, the feedforward control amount may be solved by iteration using only the gradient method described above. As still another example, the feedforward control amount may be solved by only iterating with the third function of S240.
S120: and calculating the error of the control quantity in real time according to the feedforward control quantity and the real-time control quantity obtained in the step S110.
The control error is calculated by the following formula
Figure BDA0002811496090000091
In the formula: Δ u0(k)、Δu1(k)、Δu2(k)、Δu3(k) And the transverse periodic variable pitch error, the longitudinal periodic variable pitch error, the total pitch error and the tail rotor pitch error of the k helicopter at the current moment are shown. u. of0_o(k-1)、u1_o(k-1)、u2_o(k-1)、u3_oAnd (k-1) is the real-time transverse periodic variable pitch, the real-time longitudinal periodic variable pitch, the real-time total pitch and the real-time tail pitch of the helicopter, namely the output control quantity of the flight control at the last moment k-1. Theta (k) is the pitch attitude angle of the helicopter at the current moment k, gamma (k) is the roll attitude angle of the helicopter at the current moment k, and Vz(k) And psi (k) is the climbing speed of the k helicopter at the current moment, and psi (k) is the heading angle of the k helicopter at the current moment. f. ofΔ0Representing a lateral error function.
Specifically, as one example, the lateral cyclic pitch error Δ u of the ktcopter at the current time k0(k) Longitudinal cyclic pitch error Δ u1(k) Total distance error Δ u2(k) And the tailPitch error Δ u3(k) The difference is directly made by using a PID method.
As another example, the collective pitch error Δ u of k helicopters at the current time2(k) And the error of the tail rotor pitch Deltau3(k) Directly making difference by using a PID method, and obtaining the transverse periodic variable pitch error delta u of the k helicopter at the current moment0(k) And longitudinal cyclic pitch error Deltau1(k) The following method was used for the estimation.
As shown in FIG. 3, the lateral cyclic variation error Δ u is implemented by an observer0(k) And longitudinal cyclic pitch error Deltau1(k) And (4) calculating. In the figure, i is 0 or 1. And the observer calculates according to the feedforward control quantity and the output control quantity at the last moment k-1 to obtain a corresponding control quantity error.
With reference to FIG. 3, the lateral cyclic pitch error Δ u0(k) The calculation method of (2) is as follows:
e0＝u0(k)-z01(k-1) (6)
wherein z is01(k-1)＝u0_o(k-1) (7)
z01(k)＝z01(k-1)+[β01e0+z02(k-1)]T (8)
z02(k)＝z02(k-1)+[A0z02(k-1)-B0z01(k-1)+C0u0(k-1)+z03(k-1)+β02]T (9)
z03(k)＝z03(k-1)+β03e0 (10)
Δu0(k)＝z03(k)/D0 (11)
Wherein T is a control period, z01(k) Is an estimated value of the roll angular velocity of the k helicopter at the current moment, z02(k) Is an estimated value of roll angular acceleration of the helicopter at the current moment k, z03(k) Is the first interference estimate, z, of k helicopters at the current time01(k-1) is the roll angular velocity of the helicopter at the previous moment k-1, z02(k-1) roll angular acceleration of the helicopter, z, at the previous moment k-103(k-1) is the first interference value, beta, of the helicopter at the last moment k-101、β02、β03To a specified value, A0、B0、C0Respectively, a first characteristic quantity identification parameter, D, of the helicopter0Identifying a parameter, u, for a first model of a helicopter0(k) For the lateral cyclic variation of k helicopter at the present moment, u0_o(k-1) is the output lateral cyclic variation of the helicopter at the previous moment k-1, e0An error is estimated for the roll control amount.
With reference to FIG. 3, the longitudinal cyclic pitch error Δ u1(k) The calculation method comprises the following steps:
e1＝u1(k)-z11(k-1)(12)
wherein z is11(k-1)＝u1_o(k-1) (13)
z11(k)＝z11(k-1)+[β11e0+z12(k-1)]T (14)
z12(k)＝z12(k-1)+[A1z12(k-1)-B1z11(k-1)+C1u1(k-1)+z13(k-1)+β12]T (15)
z13(k)＝z13(k-1)+β13e1 (16)
Δu1(k)＝z13(k)/D1 (17)
Wherein T is a control period, z11(k) Is an estimated value of the pitch angle speed of the k helicopter at the current moment, z12(k) Is an estimated value of the pitch angle acceleration of the k helicopter at the current moment, z13(k) Is a second interference estimate, z, for the current time k helicopter11(k-1) is the pitch angular velocity of the helicopter at the previous moment k-1, z12(k-1) is the pitch acceleration of the helicopter at the previous moment k-1, z13(k-1) is the second interference value, beta, of the helicopter at the last moment k-111、β12、β13To a specified value, A1、B1、C1Respectively, a second characteristic quantity identification parameter, D, of the helicopter1Identifying a parameter, u, for a second model of the helicopter1(k) For the lateral cyclic variation of k helicopter at the present moment, u1_o(k-1) is the last oneOutput lateral cyclic variation of a carving k-1 helicopter e1An error is estimated for the pitch control amount.
S130: and calculating the feedback control quantity according to the real-time navigation measurement data and the flight reference quantity, wherein the calculation formula is as follows.
Figure BDA0002811496090000111
Wherein, Vrefy(k) The reference side flying speed, V, of the helicopter at the current moment ky(k) Actual lateral flying speed of helicopter measured for current time k, ay(k) The actual lateral flight acceleration of the helicopter measured for the current time k, and gamma (k) the roll attitude angle, omega, of the helicopter measured for the current time kxb(k) Roll angular velocity, V, of the helicopter measured for the current time krefx(k) Reference forward flight speed, V, of helicopter at current time kx(k) Actual forward flight speed of helicopter, measured for current time k, ax(k) The actual forward flight acceleration of the helicopter measured for the current time k, and θ (k) is the pitch attitude angle of the helicopter measured for the current time k, ωyb(k) The pitch angle velocity, V, of the helicopter measured for the current time krefz(k) Is the reference climbing speed, h, of the helicopter at the current moment kref(k) K the reference altitude of the helicopter at the current time, h (k) the actual altitude of the helicopter measured at the current time k, Vz(k) Actual climbing speed of helicopter measured for current time k, az(k) Actual climb acceleration, ω, of the helicopter measured for the current time krefzb(k) Is the reference course angular velocity, psi, of the helicopter at the current time kref(k) Is the reference course angle, omega, of the helicopter at the current moment kzb(k) Actual heading angular velocity of the helicopter measured for the current time k, ψ (k) being the actual heading angle of the helicopter measured for the current time k, u0f(k)、u1f(k)、u2f(k)、u3f(k) Respectively are feedback control quantities of a roll channel, a pitch channel, a high-direction channel and a course channel of the k helicopter at the current moment. f. ofo0To represent
S140: and determining an output control quantity according to the feedforward control quantity obtained in the S110, the control quantity error obtained in the S120 and the feedback control quantity obtained in the S130, and using the output control quantity as a rudder instruction of the helicopter.
As one example, the output control amount is a combination of a feedforward control amount, a control amount error, and a feedback control amount.
As an example, the output control amount is calculated according to a summation manner:
Figure BDA0002811496090000121
wherein u is0_o(k) Represents the output transverse cyclic variation, u, of the k helicopter at the current moment1_o(k) Represents the output longitudinal cyclic variation, u, of the helicopter at the current moment k2_o(k) Represents the output collective pitch u of the helicopter at the current moment k3_o(k) Representing the output tail pitch of the helicopter at the current time k.
In the next cycle, the output control quantity at the current time k is taken as the real-time control quantity calculation control quantity error in S120, and the rudder instruction cycle output of flight control is realized.
Example two
The present application further provides a helicopter control system 400 according to the first embodiment, as shown in fig. 4, including a feedforward control amount calculation module 410, a control amount error calculation module 420, a feedback control amount calculation module 430, and a rudder instruction determination module 440;
wherein, the feedforward control quantity calculating module 410 calculates the feedforward control quantity according to the off-line pneumatic data;
the control quantity error calculation module 420 calculates a control quantity error between the feedforward control quantity and the real-time control quantity according to the real-time control quantity;
the feedback control amount calculation module 430 calculates a feedback control amount according to the real-time measurement data;
the rudder instruction determining module 440 determines an output control quantity according to the feedforward control quantity, the control quantity error and the feedback control quantity, and uses the output control quantity as a rudder instruction of the helicopter.
The beneficial effect of this application is as follows:
1. the method adopts a real-time estimation compensation control method, and combines the feedforward control quantity and the feedback control quantity obtained by real-time measurement data to improve the wind resistance control performance of the helicopter, so that the helicopter is ensured to always work near the actual trim state.
2. The method and the device utilize the balancing quantity calculated by the off-line pneumatic data to form the feedforward control quantity, reduce the channel coupling difficulty of the helicopter and improve the control response speed of the helicopter.
3. According to the method and the device, the measurement information of the sensor is fully introduced to form closed-loop feedback control, and the influence of wind disturbance on the system performance is reduced.
While the preferred embodiments of the present application have been described, additional variations and modifications in those embodiments may occur to those skilled in the art once they learn of the basic inventive concepts. Therefore, it is intended that the appended claims be interpreted as including preferred embodiments and all alterations and modifications as fall within the scope of the application. It will be apparent to those skilled in the art that various changes and modifications may be made in the present application without departing from the spirit and scope of the application. Thus, if such modifications and variations of the present application fall within the scope of the claims of the present application and their equivalents, the present application is intended to include such modifications and variations as well.

Claims (10)

1. A method of controlling a helicopter, comprising:
the following steps are executed in a circulating way:
calculating a feedforward control quantity in real time according to the off-line pneumatic data;
calculating a control quantity error between the feedforward control quantity and the real-time control quantity according to the real-time control quantity;
calculating feedback control quantity according to the real-time navigation measurement data and the flight reference quantity;
and determining an output control quantity according to the feedforward control quantity, the control quantity error and the feedback control quantity, and taking the output control quantity as a rudder instruction of the helicopter.
2. The control method according to claim 1, wherein the output control amount is a combination of the feedforward control amount, the control amount error, and the feedback control amount.
3. The control method according to claim 1 or 2, characterized in that the feedback control amount is calculated using the following formula:
Figure FDA0002811496080000011
wherein, Vrefy(k) The reference side flying speed, V, of the helicopter at the current moment ky(k) K actual side flight speed of helicopter at current moment, ay(k) The actual lateral flight acceleration of the helicopter at the current moment k, and gamma (k) is the roll attitude angle, omega, of the helicopter at the current moment kxb(k) At the current moment k, the roll angular velocity, V, of the helicopterrefx(k) Reference forward flight speed, V, of helicopter at current time kx(k) K actual forward flight speed of helicopter at current moment, ax(k) The actual forward flight acceleration of the helicopter at the current moment k, and theta (k) is the pitch attitude angle of the helicopter at the current moment k, omegayb(k) Is the pitch angle velocity, V, of the helicopter at the current moment krefz(k) Is the reference climbing speed, h, of the helicopter at the current moment kref(k) Is the reference flight height of k helicopter at the current moment, h (k) is the actual flight height of k helicopter at the current moment, Vz(k) K actual climbing speed of helicopter at present moment, az(k) Is the actual climbing acceleration, omega, of the helicopter at the current moment krefzb(k) Is the reference course angular velocity, psi, of the helicopter at the current time kref(k) Is the reference course angle, omega, of the helicopter at the current moment kzb(k) Is the actual course angular velocity of k helicopter at the current moment, psi (k) is the actual course angle of k helicopter at the current moment, u0f(k)、u1f(k)、u2f(k)、u3f(k) Respectively are feedback control quantities of a roll channel, a pitch channel, a high-direction channel and a course channel of the k helicopter at the current moment.
4. A control method according to claim 1, characterized in that the feedforward control quantity comprises at least the transverse cyclic pitch, the longitudinal cyclic pitch, the collective pitch and the tail pitch of the helicopter.
5. The control method of claim 4, wherein the feedforward control amount is calculated by solving a first function as follows based on the off-line pneumatic data
Figure FDA0002811496080000021
Wherein V is the offline flight speed of the helicopter, theta is the offline pitch attitude angle of the helicopter, gamma is the offline roll attitude angle of the helicopter, and FZbFor total off-line resultant force acting vertically on the helicopter body, MXb、MYb、MZbRespectively, the total moment of the offline, u, acting on the helicopter body in the direction X, Y, Z0For transverse cyclic pitch, u, of helicopters1For longitudinal cyclic pitch, u, of helicopters2Is the collective pitch u of the helicopter3Is the tail pitch of the helicopter.
6. The control method according to claim 5, wherein the control amount error includes a lateral cyclic pitch error Δ u0(k) Longitudinal cyclic pitch error Δ u1(k) Total distance error Δ u2(k) Tail rotor pitch error delta u3(k)；
The following formula is adopted to calculate the transverse periodic variable pitch error delta u0(k)
e0＝u0(k)-z01(k-1)
Wherein z is01(k-1)＝u0_o(k-1)
z01(k)＝z01(k-1)+[β01e0+z02(k-1)]T
z02(k)＝z02(k-1)+[A0z02(k-1) -B0z01(k-1)+C0u0(k-1)+z03(k-1)+β02]T
z03(k)＝z03(k-1)+β03e0
Δu0(k)＝z03(k)/D0
Wherein T is a control period, z01(k) Is an estimated value of the roll angular velocity of the k helicopter at the current moment, z02(k) Is an estimated value of roll angular acceleration of the helicopter at the current moment k, z03(k) Is the first interference estimate, z, of k helicopters at the current time01(k-1) is the roll angular velocity of the helicopter at the previous moment k-1, z02(k-1) roll angular acceleration of the helicopter, z, at the previous moment k-103(k-1) is the first interference value, beta, of the helicopter at the last moment k-101、β02、β03To a specified value, A0、B0、C0Respectively, a first characteristic quantity identification parameter, D, of the helicopter0Identifying a parameter, u, for a first model of a helicopter0(k) For the lateral cyclic variation of k helicopter at the present moment, u0_o(k-1) is the output lateral cyclic variation of the helicopter at the previous moment k-1, e0An error is estimated for the roll control amount.
7. The control method according to claim 6, wherein the longitudinal cyclic pitch error Δ u is calculated using the following formula1(k)
e1＝u1(k)-z11(k-1)
Wherein z is11(k-1)＝u1_o(k-1)
z11(k)＝z11(k-1)+[β11e0+z12(k-1)]T
z12(k)＝z12(k-1)+[A1z12(k-1)-B1z11(k-1)+C1u1(k-1)+z13(k-1)+β12]T
z13(k)＝z13(k-1)+β13e1
Δu1(k)＝z13(k)/D1
Wherein T is a control period, z11(k) Is an estimated value of the pitch angle speed of the k helicopter at the current moment, z12(k) Is an estimated value of the pitch angle acceleration of the k helicopter at the current moment, z13(k) Is a second interference estimate, z, for the current time k helicopter11(k-1) is the pitch angular velocity of the helicopter at the previous moment k-1, z12(k-1) is the pitch acceleration of the helicopter at the previous moment k-1, z13(k-1) is the second interference value, beta, of the helicopter at the last moment k-111、β12、β13To a specified value, A1、B1、C1Respectively, a second characteristic quantity identification parameter, D, of the helicopter1Identifying a parameter, u, for a second model of the helicopter1(k) For the lateral cyclic variation of k helicopter at the present moment, u1_o(k-1) is the output lateral cyclic variation of the helicopter at the previous moment k-1, e1An error is estimated for the pitch control amount.
8. The control method of claim 5, wherein the first function is solved by a gradient method for a second function
Figure FDA0002811496080000031
Wherein f isi 2(x) Representing the square of the target error function term.
9. The control method of claim 8, wherein the first function is solved using the following formula
x(k+1)＝x(k)-[f′(x(k))]-1f(x(k))
Figure FDA0002811496080000041
Wherein x is(k+1)Indicates the next step to be solvedParameter value, x(k)Representing the current value of the parameter to be found, f' (x)(k)) Jacobian matrix, f (x), representing the error function versus the unknown quantity(k)) A resolved value representing a first function is expressed,
Figure FDA0002811496080000042
representing the total off-line resultant force F acting on the vertical direction of the helicopter body in the first functionZbFor transverse cyclic variation u of helicopter0The partial derivative is calculated and the partial derivative is calculated,
Figure FDA0002811496080000043
representing the total off-line resultant force F acting on the vertical direction of the helicopter body in the first functionZbFor longitudinal cyclic variation u of helicopter1The partial derivative is calculated and the partial derivative is calculated,
Figure FDA0002811496080000044
representing the total off-line resultant force F acting on the vertical direction of the helicopter body in the first functionZbTotal distance u to helicopter2The partial derivative is calculated and the partial derivative is calculated,
Figure FDA0002811496080000045
representing the total off-line resultant force F acting on the vertical direction of the helicopter body in the first functionZbTail rotor pitch u for helicopter3Calculating a partial derivative;
Figure FDA0002811496080000046
representing the sum of the moments M taken off-line in the first function and acting in the X direction on the helicopter bodyXbFor transverse cyclic variation u of helicopter0The partial derivative is calculated and the partial derivative is calculated,
Figure FDA0002811496080000047
representing the sum of the moments M taken off-line in the first function and acting in the X direction on the helicopter bodyXbFor longitudinal cyclic variation u of helicopter1The partial derivative is calculated and the partial derivative is calculated,
Figure FDA0002811496080000048
representing the sum of the moments M taken off-line in the first function and acting in the X direction on the helicopter bodyXbTotal distance u to helicopter2The partial derivative is calculated and the partial derivative is calculated,
Figure FDA0002811496080000049
representing the sum of the moments M taken off-line in the first function and acting in the X direction on the helicopter bodyXbTail rotor pitch u for helicopter3Calculating a partial derivative;
Figure FDA00028114960800000410
representing the sum of the moments M taken off-line in the first function, acting in the direction Y on the helicopter bodyYbFor transverse cyclic variation u of helicopter0The partial derivative is calculated and the partial derivative is calculated,
Figure FDA0002811496080000051
representing the sum of the moments M taken off-line in the first function, acting in the direction Y on the helicopter bodyYbLongitudinal cyclic variable pitch u of medium-to-vertical lift1The partial derivative is calculated and the partial derivative is calculated,
Figure FDA0002811496080000052
representing the sum of the moments M taken off-line in the first function, acting in the direction Y on the helicopter bodyXbTotal distance u to helicopter2The partial derivative is calculated and the partial derivative is calculated,
Figure FDA0002811496080000053
representing the sum of the moments M taken off-line in the first function, acting in the direction Y on the helicopter bodyXbTail rotor pitch u for helicopter3Calculating a partial derivative;
Figure FDA0002811496080000054
representing the total moment M of the first function in the Z direction acting on the helicopter bodyZbTransverse to the helicopterPeriodic variable pitch u0The partial derivative is calculated and the partial derivative is calculated,
Figure FDA0002811496080000055
representing the total moment M of the first function in the Z direction acting on the helicopter bodyZbLongitudinal cyclic variable pitch u of medium-to-vertical lift1The partial derivative is calculated and the partial derivative is calculated,
Figure FDA0002811496080000056
representing the total moment M of the first function in the Z direction acting on the helicopter bodyZbTotal distance u to helicopter2The partial derivative is calculated and the partial derivative is calculated,
Figure FDA0002811496080000057
representing the total moment M of the first function in the Z direction acting on the helicopter bodyZbTail rotor pitch u for helicopter3And (5) calculating a partial derivative.
10. A control system of a helicopter is characterized by comprising a feedforward control quantity calculation module, a control quantity error calculation module, a feedback control quantity calculation module and a rudder instruction determination module;
the feedforward control quantity calculation module calculates the feedforward control quantity in real time according to the off-line pneumatic data;
the control quantity error calculation module calculates a control quantity error between the feedforward control quantity and the real-time control quantity according to the real-time control quantity;
the feedback control quantity calculation module calculates the feedback control quantity according to the real-time navigation measurement data and the flight reference quantity;
and the rudder instruction determining module determines an output control quantity according to the feedforward control quantity, the control quantity error and the feedback control quantity, and takes the output control quantity as a rudder instruction of the helicopter.
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