CN112486218B - Helicopter control method and system - Google Patents
Helicopter control method and system Download PDFInfo
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- CN112486218B CN112486218B CN202011387658.4A CN202011387658A CN112486218B CN 112486218 B CN112486218 B CN 112486218B CN 202011387658 A CN202011387658 A CN 202011387658A CN 112486218 B CN112486218 B CN 112486218B
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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. According to the method, a real-time estimation compensation control method is adopted, and the wind resistance control performance of the helicopter is improved by combining the feedforward control quantity and the feedback control quantity obtained by real-time navigation measurement data, so that the helicopter is ensured to work nearby an actual balancing state all the time.
Description
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:
wherein, V refy (k) The reference side flying speed, V, of the helicopter at the current moment k y (k) K actual side flight speed of helicopter at current moment, a y (k) The actual side flying acceleration of the helicopter at the current moment k, gamma (k) is the roll attitude angle of the helicopter at the current moment k, and omega xb (k) At the current moment k, the roll angular velocity, V, of the helicopter refx (k) Reference forward flight speed, V, of helicopter at current time k x (k) K actual forward flight speed of helicopter at current moment, a x (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, omega yb (k) Is the pitch angle velocity, V, of the helicopter at the current moment k refz (k) Is the reference climbing speed, h, of the helicopter at the current moment k ref (k) Is the reference flight height of k helicopters at the current moment, h (k) is the actual flight height of k helicopters at the current moment, V z (k) At the current moment k, the actual climbing speed of the helicopter, a z (k) Is the actual climbing acceleration, omega, of the helicopter at the current moment k refzb (k) Is the reference course angular velocity, psi, of the helicopter at the current time k ref (k) Is the reference course angle, omega, of the helicopter at the current moment k zb (k) Is the actual heading angular velocity of k helicopter at the current moment, psi (k) is the actual heading angle of k helicopter at the current moment, u 0f (k)、u 1f (k)、u 2f (k)、u 3f (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
Wherein V is the offline flying speed of the helicopter, theta is the offline pitching attitude angle of the helicopter, and gamma is the distance of the helicopterLine roll attitude angle, F Zb For total off-line resultant force acting vertically on the helicopter body, M Xb 、M Yb 、M Zb Respectively, the offline total moment u acting on the helicopter body in the direction of X, Y, Z 0 For transverse cyclic pitch, u, of helicopters 1 For longitudinal cyclic pitch, u, of helicopters 2 Is the collective pitch u of the helicopter 3 Is the tail pitch of the helicopter.
Preferably, the control quantity error includes a lateral cyclic pitch error Δ u 0 (k) Longitudinal cyclic pitch error Δ u 1 (k) Total distance error Δ u 2 (k) Tail rotor pitch error delta u 3 (k)；
The following formula is adopted to calculate the transverse periodic variable pitch error delta u 0 (k)
e 0 ＝u 0 (k)-z 01 (k-1)
Wherein z is 01 (k-1)＝u 0_o (k-1)
z 01 (k)＝z 01 (k-1)+[β 01 e 0 +z 02 (k-1)]T
z 02 (k)＝z 02 (k-1)+[A 0 z 02 (k-1)-B 0 z 01 (k-1)+C 0 u 0 (k-1)+z 03 (k-1)+β 02 ]T
z 03 (k)＝z 03 (k-1)+β 03 e 0
Δu 0 (k)＝z 03 (k)/D 0
Wherein T is a control period, z 01 (k) Is an estimated value of the roll angular velocity of the k helicopter at the current moment, z 02 (k) Is an estimated value of roll angular acceleration of the helicopter at the current moment k, z 03 (k) Is the first interference estimate, z, of k helicopters at the current time 01 (k-1) is the roll angular velocity of the helicopter at the previous moment k-1, z 02 (k-1) roll angular acceleration of the helicopter, z, at the previous moment k-1 03 (k-1) is a first disturbance value, beta, of the helicopter at the previous moment k-1 01 、β 02 、β 03 To a specified value, A 0 、B 0 、C 0 Are respectively a helicopterFirst characteristic quantity identification parameter of machine, D 0 Identifying a parameter, u, for a first model of a helicopter 0 (k) For the lateral cyclic variation of k helicopter at the present moment, u 0_o (k-1) is the output lateral cyclic variation of the helicopter at the previous moment k-1, e 0 An error is estimated for the roll control amount.
Preferably, the longitudinal cyclic pitch error Δ u is calculated using the following formula 1 (k)
e 1 ＝u 1 (k)-z 11 (k-1)
Wherein z is 11 (k-1)＝u 1_o (k-1)
z 11 (k)＝z 11 (k-1)+[β 11 e 0 +z 12 (k-1)]T
z 12 (k)＝z 12 (k-1)+[A 1 z 12 (k-1)-B 1 z 11 (k-1)+C 1 u 1 (k-1)+z 13 (k-1)+β 12 ]T
z 13 (k)＝z 13 (k-1)+β 13 e 1
Δu 1 (k)＝z 13 (k)/D 1
Wherein T is a control period, z 11 (k) Is an estimated value of the pitch angle speed of the k helicopter at the current moment, z 12 (k) Is an estimated value of the pitch angle acceleration of the k helicopter at the current moment, z 13 (k) Is a second interference estimate, z, for the current time k helicopter 11 (k-1) is the pitch angular velocity of the helicopter at the previous moment k-1, z 12 (k-1) is the pitch acceleration of the helicopter at the previous moment k-1, z 13 (k-1) is a second disturbance value, beta, of the helicopter at the previous time k-1 11 、β 12 、β 13 To a specified value, A 1 、B 1 、C 1 Respectively, a second characteristic quantity identification parameter, D, of the helicopter 1 Identifying a parameter, u, for a second model of the helicopter 1 (k) For the lateral cyclic variation of k helicopter at the present moment, u 1_o (k-1) is the output lateral cyclic variation of the helicopter at the previous moment k-1, e 1 An error is estimated for the pitch control amount.
Preferably, the first function is solved by using a gradient method for the following second function
Wherein f is i 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) )] -1 f(x (k) )
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) ) A resolved value representing a first function is expressed,
showing the total off-line resultant F of the first function acting on the vertical direction of the helicopter body Zb For transverse cyclic variation u of helicopter 0 The partial derivative is calculated and the partial derivative is calculated,
representing the total off-line resultant force F acting on the vertical direction of the helicopter body in the first function Zb For longitudinal cyclic variation u of helicopter 1 The partial derivative is calculated and the partial derivative is calculated,
representing the total off-line resultant force F acting on the vertical direction of the helicopter body in the first function Zb Total distance u to helicopter 2 The partial derivative is calculated and the partial derivative is calculated,
means for indicating whether the first function is acting in the vertical direction of the helicopter bodyResultant force F Zb Tail rotor pitch u for helicopter 3 Calculating a partial derivative;
representing the sum of the moments M taken off-line in the first function and acting in the X direction on the helicopter body Xb For transverse cyclic variation u of helicopter 0 The partial derivative is calculated and the partial derivative is calculated,
representing the sum of the moments M taken off-line in the first function and acting in the X direction on the helicopter body Xb For longitudinal cyclic variation u of helicopter 1 The partial derivative is calculated and the partial derivative is calculated,
representing the sum of the moments M taken off-line in the first function and acting in the X direction on the helicopter body Xb Total distance u to helicopter 2 The partial derivative is calculated and the partial derivative is calculated,
representing the sum of the moments M taken off-line in the first function and acting in the X direction on the helicopter body Xb Tail rotor pitch u for helicopter 3 Calculating a partial derivative;
representing the sum of the moments M taken off-line in the first function, acting in the direction Y on the helicopter body Yb For transverse cyclic variation u of helicopter 0 The partial derivative is calculated and the partial derivative is calculated,
representing the sum of the moments M taken off-line in the first function, acting in the direction Y on the helicopter body Yb Longitudinal cyclic variable pitch u of medium-to-vertical lift 1 The partial derivative is calculated and the partial derivative is calculated,
representing the sum of the moments M taken off-line in the first function, acting in the direction Y on the helicopter body Xb Total distance u to helicopter 2 The partial derivative is calculated and the partial derivative is calculated,
representing the sum of the moments M taken off-line in the first function, acting in the direction Y on the helicopter body Xb Tail rotor pitch u for helicopter 3 Calculating a partial derivative;
representing the total moment M of the first function in the Z direction acting on the helicopter body Zb For transverse cyclic variation u of helicopter 0 The partial derivative is calculated and the partial derivative is calculated,
representing the total moment M of the first function in the Z direction acting on the helicopter body Zb Longitudinal cyclic variable pitch u of medium-to-vertical lift 1 The partial derivative is calculated and the partial derivative is calculated,
representing the total moment M of the first function in the Z direction acting on the helicopter body Zb Total distance u to helicopter 2 The partial derivative is calculated and the partial derivative is calculated,
representing the total moment M of the first function in the Z direction acting on the helicopter body Zb Tail rotor pitch u for helicopter 3 And (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 used in the description of the embodiments or the prior art will be briefly described below, it is obvious that the drawings in the description below are only some embodiments described in the present application, and other drawings can be obtained by those skilled in the art according to these 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
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 F Zb For total off-line resultant force acting vertically on the helicopter body, M Xb 、M Yb 、M Zb Respectively, the off-line total moment u acting on the helicopter body in the direction of X, Y, Z 0 For transverse cyclic pitch, u, of helicopters 1 For longitudinal cyclic pitch, u, of helicopters 2 Is the collective pitch of the helicopter, u 3 Is the tail pitch of the helicopter. Wherein the feedforward control quantity comprises u 0 、u 1 、u 2 、u 3 As unknown. V, theta, gamma, F Zb ，M Xb ，M Yb ，M Zb In 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
Wherein f is i 2 (x) Representing the square of the target error function term.
From the initial point
Starting along a second function
Is the fastest decreasing partyTo (i.e. in the direction of the negative gradient), gradually decrease
The value is obtained.
S220: judging whether the latest solution value is less than the first extreme value epsilon 1 (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) )] -1 f(x (k) ) (3)
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) ) A resolved value of the first function is represented,
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,
representing F in a first function Zb For u is paired 0 The partial derivative is calculated and the partial derivative is calculated,
representing F in a first function Zb For u is paired 1 The partial derivative is calculated and the partial derivative is calculated,
representing F in a first function Zb The partial derivative is calculated for u2,
representing F in a first function Zb For u is paired 3 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 ε 2 (ii) a If yes, go to S260; otherwise, return to step S240. Wherein, Δ u i (k) Represents the ith (i is belonged to {0,1,2,3 }) feedforward control quantity u at the current time k i (k) The error of (a) is detected,
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 output 0 、u 1 、u 2 、u 3 Feedforward lateral cyclic variation u of helicopter as current moment k 0 (k) Feedforward longitudinal cyclic variation u 1 (k) Feedforward total distance u 2 (k) And feed forward tail pitch u 3 (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
In the formula: Δ u 0 (k)、Δu 1 (k)、Δu 2 (k)、Δu 3 (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. u 0_o (k-1)、u 1_o (k-1)、u 2_o (k-1)、u 3_o And (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 V z (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. of Δ0 Representing a lateral error function.
Specifically, as one example, the lateral cyclic pitch error Δ u of the ktcopter at the current time k 0 (k) Longitudinal cyclic pitch error Δ u 1 (k) Total distance error Δ u 2 (k) And the error of the tail rotor pitch Deltau 3 (k) And directly performing difference by using a PID method.
As another example, the collective pitch error Δ u of helicopter at the current time k 2 (k) And the tail pitch error Deltau 3 (k) Directly performing difference by using a PID method, and determining the transverse periodic variable pitch error delta u of the k helicopter at the current moment 0 (k) And longitudinal cyclic pitch error Deltau 1 (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 observer 0 (k) And longitudinal cyclic pitch error Deltau 1 (k) And (4) calculating. In the figure, i =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 u 0 (k) The calculation method of (2) is as follows:
e 0 ＝u 0 (k)-z 01 (k-1) (6)
wherein z is 01 (k-1)＝u 0_o (k-1) (7)
z 01 (k)＝z 01 (k-1)+[β 01 e 0 +z 02 (k-1)]T (8)
z 02 (k)＝z 02 (k-1)+[A 0 z 02 (k-1)-B 0 z 01 (k-1)+C 0 u 0 (k-1)+z 03 (k-1)+β 02 ]T (9)
z 03 (k)＝z 03 (k-1)+β 03 e 0 (10)
Δu 0 (k)＝z 03 (k)/D 0 (11)
Wherein T is a control period, z 01 (k) Is an estimated value of the roll angular velocity of the k helicopter at the current moment, z 02 (k) Is an estimated value of roll angular acceleration of the helicopter at the current moment k, z 03 (k) Is the first interference estimate, z, of k helicopters at the current time 01 (k-1) is the roll angular velocity of the helicopter at the previous time k-1, z 02 (k-1) roll angular acceleration of the helicopter, z, at the previous moment k-1 03 (k-1) is the first interference value, beta, of the helicopter at the last moment k-1 01 、β 02 、β 03 To a specified value, A 0 、B 0 、C 0 Respectively, a first characteristic quantity identification parameter, D, of the helicopter 0 Identifying a parameter, u, for a first model of a helicopter 0 (k) For the lateral cyclic variation of k helicopter at the present moment, u 0_o (k-1) is the output lateral cyclic variation of the helicopter at the previous moment k-1, e 0 An error is estimated for the roll control amount.
With reference to FIG. 3, the longitudinal cyclic pitch error Δ u 1 (k) The calculation method comprises the following steps:
e 1 ＝u 1 (k)-z 11 (k-1)(12)
wherein z is 11 (k-1)＝u 1_o (k-1) (13)
z 11 (k)＝z 11 (k-1)+[β 11 e 0 +z 12 (k-1)]T (14)
z 12 (k)＝z 12 (k-1)+[A 1 z 12 (k-1)-B 1 z 11 (k-1)+C 1 u 1 (k-1)+z 13 (k-1)+β 12 ]T (15)
z 13 (k)＝z 13 (k-1)+β 13 e 1 (16)
Δu 1 (k)＝z 13 (k)/D 1 (17)
Wherein T is a control period, z 11 (k) Is an estimated value of the pitch angle speed of the k helicopter at the current moment, z 12 (k) Is an estimated value of the pitch angle acceleration of the k helicopter at the current moment, z 13 (k) Is a second interference estimate, z, for the current time k helicopter 11 (k-1) is the pitch angular velocity of the helicopter at the previous moment k-1, z 12 (k-1) is the pitch acceleration of the helicopter at the previous moment k-1, z 13 (k-1) is the second interference value, beta, of the helicopter at the last moment k-1 11 、β 12 、β 13 To a specified value, A 1 、B 1 、C 1 Respectively, a second characteristic quantity identification parameter, D, of the helicopter 1 Identifying a parameter, u, for a second model of the helicopter 1 (k) For the lateral cyclic variation of k helicopter at the present moment, u 1_o (k-1) is the output transverse cyclic variation of the helicopter at the previous moment k-1, e 1 An 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.
Wherein, V refy (k) The reference side flying speed, V, of the helicopter at the current moment k y (k) Actual lateral flying speed of helicopter measured for current time k, a y (k) The actual side-flying acceleration of the helicopter is measured at the current moment k, and gamma (k) is the roll attitude angle, omega, of the helicopter is measured at the current moment k xb (k) Roll angular velocity, V, of the helicopter measured for the current time k refx (k) Is the reference forward flight speed, V, of the helicopter at the current time k x (k) Actual forward flight speed of helicopter measured for current time k, a x (k) Measured for the current time kThe actual forward flight acceleration of the helicopter, theta (k), is the pitch attitude angle of the helicopter measured at the current time k, omega yb (k) The pitch angle velocity, V, of the helicopter measured for the current time k refz (k) Is the reference climbing speed, h, of the helicopter at the current moment k ref (k) K is the reference flight height of the helicopter at the current moment, h (k) is the actual flight height of the helicopter measured at the current moment k, V z (k) Actual climbing speed of helicopter measured for current time k, a z (k) Actual climb acceleration, ω, of the helicopter measured for the current time k refzb (k) Is the reference course angular velocity psi of the helicopter at the current time k ref (k) Is the reference course angle, omega, of the helicopter at the current moment k zb (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, u 0f (k)、u 1f (k)、u 2f (k)、u 3f (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. of o0 To 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:
wherein u is 0_o (k) Represents the output transverse cyclic variation, u, of the k helicopter at the current moment 1_o (k) Represents the output longitudinal cyclic variation, u, of the helicopter at the current moment k 2_o (k) Represents the output collective pitch u of the helicopter at the current moment k 3_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 (7)
1. A method of controlling a helicopter, comprising:
the following steps are executed in a circulating way:
calculating 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;
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;
wherein the feedback control amount is calculated using the following formula:
wherein, V refy (k) The reference side flying speed, V, of the helicopter at the current moment k y (k) K actual side flight speed of helicopter at current moment, a y (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 k xb (k) At the current moment k, the roll angular velocity, V, of the helicopter refx (k) Reference forward flight speed, V, of helicopter at current time k x (k) K actual forward flight speed of helicopter at current moment, a x (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, omega yb (k) Is the pitch angle velocity, V, of the helicopter at the current moment k refz (k) Is the reference climbing speed, h, of the helicopter at the current moment k ref (k) For k rise at the current momentReference flight height of the helicopter, h (k) is the actual flight height of the helicopter at the current moment k, V z (k) At the current moment k, the actual climbing speed of the helicopter, a z (k) Is the actual climbing acceleration, omega, of the helicopter at the current moment k refzb (k) Is the reference course angular velocity, psi, of the helicopter at the current time k ref (k) Is the reference course angle, omega, of the helicopter at the current moment k zb (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, u 0f (k)、u 1f (k)、u 2f (k)、u 3f (k) Respectively feeding back 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;
calculating feedforward control quantity by solving the following first function according to off-line pneumatic data
Wherein V is the offline flying speed of the helicopter, theta is the offline pitching attitude angle of the helicopter, gamma is the offline rolling attitude angle of the helicopter, and F Zb For total off-line resultant force acting vertically on the helicopter body, M Xb 、M Yb 、M Zb Respectively, the offline total moment u acting on the helicopter body in the direction of X, Y, Z 0 For transverse cyclic pitch, u, of helicopters 1 For longitudinal cyclic pitch, u, of helicopters 2 Is the collective pitch u of the helicopter 3 Is the tail pitch of the helicopter.
2. 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.
3. The control method according to claim 2, wherein the control amount error includes a lateral cyclic pitch error Δ u 0 (k) Longitudinal cyclic pitch error Δ u 1 (k) Total distance error Δ u 2 (k) Tail rotor pitch error delta u 3 (k)；
The following formula is adopted to calculate the transverse periodic variable pitch error delta u 0 (k)
e 0 ＝u 0 (k)-z 01 (k-1)
Wherein z is 01 (k-1)＝u 0_o (k-1)
z 01 (k)＝z 01 (k-1)+[β 01 e 0 +z 02 (k-1)]T
z 02 (k)＝z 02 (k-1)+[A 0 z 02 (k-1)-B 0 z 01 (k-1)+C 0 u 0 (k-1)+z 03 (k-1)+β 02 ]T
z 03 (k)＝z 03 (k-1)+β 03 e 0
Δu 0 (k)＝z 03 (k)/D 0
Wherein T is a control period, z 01 (k) Is an estimated value of the roll angular velocity of the k helicopter at the current moment, z 02 (k) Is an estimated value of roll angular acceleration of the helicopter at the current moment k, z 03 (k) Is the first interference estimate, z, of k helicopters at the current time 01 (k-1) is the roll angular velocity of the helicopter at the previous moment k-1, z 02 (k-1) roll angular acceleration of the helicopter, z, at the previous moment k-1 03 (k-1) is a first disturbance value, beta, of the helicopter at the previous moment k-1 01 、β 02 、β 03 To a specified value, A 0 、B 0 、C 0 Respectively, a first characteristic quantity identification parameter, D, of the helicopter 0 Identifying a parameter, u, for a first model of a helicopter 0 (k) For the lateral cyclic variation of k helicopter at the present moment, u 0_o (k-1) is the output lateral cyclic variation of the helicopter at the previous moment k-1, e 0 Estimating an error, u, for roll control quantities 0 (k-1) represents the lateral cyclic variation of the helicopter at the last time k-1.
4. A control method according to claim 3, characterized in that the longitudinal cyclic pitch error au is calculated using the following formula 1 (k)
e 1 ＝u 1 (k)-z 11 (k-1)
Wherein z is 11 (k-1)＝i 1_o (k-1)
z 11 (k)＝z 11 (k-1)+[β 11 e 1 +z 12 (k-1)]T
z 12 (k)＝z 12 (k-1)+[A 1 z 12 (k-1)-B 1 z 11 (k-1)+C 1 u 1 (k-1)+z 13 (k-1)+β 12 ]T
z 13 (k)＝z 13 (k-1)+β 13 e 1
Δu 1 (k)＝z 13 (k)/D 1
Wherein T is a control period, z 11 (k) Is an estimated value of the pitch angle speed of the k helicopter at the current moment, z 12 (k) Is an estimated value of the pitch angle acceleration of the k helicopter at the current moment, z 13 (k) Is a second interference estimate, z, for the helicopter at the current time k 11 (k-1) is the pitch angular velocity of the helicopter at the previous moment k-1, z 12 (k-1) is the pitch acceleration of the helicopter at the previous moment k-1, z 13 (k-1) is the second interference value, beta, of the helicopter at the last moment k-1 11 、β 12 、β 13 To a given value, A 1 、B 1 、C 1 Respectively, a second characteristic quantity identification parameter, D, of the helicopter 1 Identifying a parameter, u, for a second model of the helicopter 1 (k) For the current time k longitudinal cyclic pitch, u, of the helicopter 1_o (k-1) is the output lateral cyclic variation of the helicopter at the previous moment k-1, e 1 Estimating an error for a pitch control quantity, u 1 (k-1) represents the longitudinal cyclic variation of the helicopter at the last moment k-1.
5. The control method of claim 1, wherein the first function is solved by a gradient method for a second function
6. The control method of claim 5, wherein the first function is solved using the following formula
x (k+1) ＝x (k) -[f′(x (k) )] -1 f(x (k) )
Wherein x is (k+1) Represents 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) ) A resolved value representing a first function is expressed,
showing the total off-line resultant F of the first function acting on the vertical direction of the helicopter body Zb For transverse cyclic variation u of helicopter 0 The partial derivative is calculated and the partial derivative is calculated,
representing the total off-line resultant force F acting on the vertical direction of the helicopter body in the first function Zb For longitudinal cyclic variation u of helicopter 1 The partial derivative is calculated and the partial derivative is calculated,
showing the total off-line resultant F of the first function acting on the vertical direction of the helicopter body Zb Total distance u to helicopter 2 The partial derivative is calculated and the partial derivative is calculated,
representing the total off-line resultant force F acting on the vertical direction of the helicopter body in the first function Zb Tail rotor pitch u for helicopter 3 Calculating a partial derivative;
7. 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;
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 the output control quantity is used as a rudder instruction of the helicopter;
wherein the feedback control amount is calculated using the following formula:
wherein, V refy (k) The reference side flying speed, V, of the helicopter at the current moment k y (k) K actual side flight speed of helicopter at current moment, a y (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 k xb (k) At the current moment k, the roll angular velocity, V, of the helicopter refx (k) Reference forward flight speed, V, of helicopter at current time k x (k) At the current moment k, the actual forward flight speed of the helicopter, a x (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, omega yb (k) Is the pitch angle velocity, V, of the helicopter at the current moment k refz (k) Is the reference climbing speed, h, of the helicopter at the current moment k ref (k) Is the reference flight height of k helicopters at the current moment, h (k) is the actual flight height of k helicopters at the current moment, V z (k) At the current moment k, the actual climbing speed of the helicopter, a z (k) Is the actual climbing acceleration, omega, of the helicopter at the current moment k refzb (k) Is the reference course angular velocity, psi, of the helicopter at the current time k ref (k) Is the reference course angle, omega, of the helicopter at the current moment k zb (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, u 0f (k)、u 1f (k)、u 2f (k)、u 3f (k) Respectively at the current moment kFeedback control quantities of a roll channel, a pitch channel, a high-direction channel and a course channel of the elevator;
calculating feedforward control quantity by solving the following first function according to off-line pneumatic data
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 F Zb For total off-line resultant force acting vertically on the helicopter body, M Xb 、M Yb 、M Zb Respectively, the offline total moment u acting on the helicopter body in the direction of X, Y, Z 0 For transverse cyclic pitch, u, of helicopters 1 For longitudinal cyclic pitch, u, of helicopters 2 Is the collective pitch u of the helicopter 3 Is the tail pitch of the helicopter.
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