CN109581878A - Control method based on the tiltrotor aircraft mode conversion stage - Google Patents
Control method based on the tiltrotor aircraft mode conversion stage Download PDFInfo
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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05B—CONTROL OR REGULATING SYSTEMS IN GENERAL; FUNCTIONAL ELEMENTS OF SUCH SYSTEMS; MONITORING OR TESTING ARRANGEMENTS FOR SUCH SYSTEMS OR ELEMENTS
- G05B13/00—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion
- G05B13/02—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric
- G05B13/04—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric involving the use of models or simulators
- G05B13/042—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric involving the use of models or simulators in which a parameter or coefficient is automatically adjusted to optimise the performance
Abstract
The present invention discloses a kind of control method based on the tiltrotor aircraft mode conversion stage, initially set up longitudinal nonlinear model of tiltrotor aircraft, according to the track following information of tiltrotor aircraft, establish longitudinal Nonlinear Error Models of tiltrotor aircraft, and by choosing suitable state variable and parameter, longitudinal Nonlinear Error Models of tiltrotor aircraft are converted into the form of nonlinear time-varying state space；Secondly, the state controller of construction nonlinear time-varying；Finally the convex optimization problem of quadratic sum is converted by the design problem of controller to solve.Control problem of the effective solution tiltrotor aircraft of the present invention in the mode conversion stage.
Description
Technical field
The present invention relates to field of aerospace technology, and in particular to one kind is based on the tiltrotor aircraft mode conversion stage
Control method.
Background technique
The tiltrotor aircraft set characteristic of fixed wing aircraft and helicopter has critically important civilian and military valence
Value.However, existing simultaneously aerodynamic arrangement and pneumatic rudder face in the manipulating mechanism of tiltrotor aircraft, tilting rotor is caused to fly
The modeling and flight control of device all have certain difficulty.Tiltrotor aircraft system have time-varying, non-linear, non-radioactive,
The dynamic characteristics such as close coupling, control redundancy.Therefore, to the research of the control problem of tiltrotor aircraft be one have choose
The work of war property, wherein the control in the mode conversion stage of tiltrotor aircraft is the most difficult and complicated.
Currently, the document in relation to tiltrotor aircraft, is substantially all and inputs using tilt angle as control.But it is modeling
All there is coupling terms, design of this modeling method to controller in other control input variables such as middle tilt angle and throttle, rudder face
Cause very big difficulty.Later, some scholars are in modeling, it is contemplated that nacelle inclining rotary mechanism can eliminate the phase of control input
Multiply item.But this processing method is primarily used to the Model control of the helicopter of processing tiltrotor aircraft, especially needs
Wanting system is linear time-varying, and all there is certain limitations for the method for above-mentioned processing.
Summary of the invention
To be solved by this invention is the control problem in tiltrotor aircraft mode conversion stage, provides one kind and is based on inclining
Rotate the control method in rotor aircraft mode conversion stage.
To solve the above problems, the present invention is achieved by the following technical solutions:
Based on the control method in tiltrotor aircraft mode conversion stage, specifically include that steps are as follows:
Step 1, the longitudinal nonlinear model for establishing tiltrotor aircraft；
The reference locus of step 2, given tiltrotor aircraft, i.e., the speed of the tiltrotor aircraft of given reference,
The angle of attack, pitch angle, pitch rate and height, and given reference along the power of body x-axis, the power of y-axis and pitching moment, will
Longitudinal nonlinear model of the tiltrotor aircraft of foundation converts error model；
Step 3 chooses state variable x and control input variable u, by longitudinal nonlinearity erron mould of tiltrotor aircraft
Type is converted into the form of easy-to-handle state space；Wherein
Step 4, by assuming that the speed of tiltrotor aircraft, the angle of attack, pitch angle, pitch rate and height can be surveyed and
It constructs nonlinear time-varying and controls controller；
Step 5, to convert quadratic sum for the design problem of the nonlinear time-varying of tiltrotor aircraft control controller convex
Optimization problem is solved, i.e.,
In the above formulas, V and V*Respectively indicate the reality and reference velocity of tiltrotor aircraft；α and α*It respectively indicates and inclines
Rotate the reality of rotor aircraft and with reference to the angle of attack；WithRespectively indicate the reality of tiltrotor aircraft and with reference to pitch angle；q
And q*Respectively indicate the reality of tiltrotor aircraft and with reference to pitch rate；H and H*Respectively indicate tiltrotor aircraft
Reality and reference altitude；FxtWithIt respectively indicates along the reality of body x-axis and with reference to force component；FytWithRespectively indicate edge
The reality of body y-axis and refer to force component；MzWithIt respectively indicates along the reality of body and with reference to pitching moment；
A12=-Fyt*/ m, A52=-V*, A53=V*；Aj(x, τ) indicates A
The jth row of (x, τ), j=1,2 ..., 5；
B11=1/m, B12=-(α*+ Δ α)/m, B43=1/Iz,
A=8.791 × 10-4, b=-0.03274, c=0.3491；
For Lyapunov matrix to be asked；The increasing of controller is controlled for nonlinear time-varying to be asked
Beneficial matrix；X indicates state variable,Indicate the state variable so that the row vector all 0 of B (x, τ), τ indicates that tilting rotor flies
The tilt angle of row device,Indicate derivative of the tilt angle to the time of tiltrotor aircraft；Withvi It respectively indicatesThe upper bound and
Lower bound；I indicates the unit matrix of suitable dimension；z1Indicate suitable dimension non-zero column vector；ε1, ε2And ε3Respectively indicate given normal number；
Step 6, using solve come nonlinear time-varying control controller tiltrotor aircraft mode can be realized
The control target in conversion stage.
In above-mentioned steps 4, the nonlinear time-varying state controller that is constructed.
In above-mentioned steps 5, the convex optimization problem of quadratic sum is solved using the tool box SOStools in Matlab.
Compared with prior art, the present invention has a characteristic that
1, the tilt angle τ of tiltrotor aircraft is regarded as parameter, and non-controlling inputs, and solves the coupling that modeling generates
Item is closed, calculation amount is reduced；
2, nonlinear time-varying controller is solved by quadratic sum convex optimized algorithm, for the derivative of parameter, is answered
The S- process, is effectively reduced conservative；
3, designed nonlinear time-varying controller is the polynomial function of state and parameter, is easy to Project Realization；
4, each reference signal of tilting rotor is effectively and rapidly traced into, there is good control effect.
Detailed description of the invention
Fig. 1 is the flow chart of the control method based on the tiltrotor aircraft mode conversion stage.
Fig. 2 is the simulation track figure of tilt angle.
Fig. 3 is the simulation track figure of angular speed of verting.
Fig. 4 is the simulation track figure of angular acceleration of verting.
Fig. 5 is the speed V pursuit path curve graph of tiltrotor aircraft.(a) represents reference velocity track in figure, (b)
The speed trajectory that linear time-varying model obtains is represented, the speed trajectory that the present invention obtains (c) is represented.
Fig. 6 is the angle of attack pursuit path curve graph of tiltrotor aircraft, and (a) represents reference velocity track in figure, (b)
The speed trajectory that linear time-varying model obtains is represented, the speed trajectory that the present invention obtains (c) is represented.
Fig. 7 is the pitch angle of tiltrotor aircraftPursuit path curve graph, (a) represents reference velocity track in figure,
(b) speed trajectory that linear time-varying model obtains is represented, the speed trajectory that the present invention obtains (c) is represented.
Fig. 8 is the pitch rate q pursuit path curve graph of tiltrotor aircraft, and (a) represents reference velocity rail in figure
Mark (b) represents the speed trajectory that linear time-varying model obtains, and (c) represents the speed trajectory that the present invention obtains.
Fig. 9 is the height H pursuit path curve graph of tiltrotor aircraft, and (a) represents reference velocity track in figure, (b)
The speed trajectory that linear time-varying model obtains is represented, the speed trajectory that the present invention obtains (c) is represented.
Figure 10 is power F of the tiltrotor aircraft along x-axisxtGeometric locus figure, (a) represents reference velocity track in figure, (b)
The speed trajectory that linear time-varying model obtains is represented, the speed trajectory that the present invention obtains (c) is represented.
Figure 11 is power F of the tiltrotor aircraft along y-axisytGeometric locus figure, (a) represents reference velocity track in figure, (b)
The speed trajectory that linear time-varying model obtains is represented, the speed trajectory that the present invention obtains (c) is represented.
Figure 12 is the pitching moment M of tiltrotor aircraftzPursuit path curve graph, (a) represents reference velocity rail in figure
Mark (b) represents the speed trajectory that linear time-varying model obtains, and (c) represents the speed trajectory that the present invention obtains.
Specific embodiment
To make the objectives, technical solutions, and advantages of the present invention clearer, below in conjunction with specific example, and referring to attached
Figure, the present invention is described in more detail.
The present invention in the modelling phase by avoiding and other controls using the tilt angle of tiltrotor aircraft as parameter
Coupling terms between vector processed carry out the design of nonlinear time-varying state controller using the information of tilt angle.However tilting rotor
Aircraft is a kind of with non-linear, the system of strong time-varying characteristics, and the calculating of corresponding Solvability condition is a problem.This hair
It is bright to utilize a kind of technology --- the quadratic sum technology risen and graduallyd mature in recent years, it will be turned based on tiltrotor aircraft mode
The stability problem for changing the stage is converted into the convex optimization problem of quadratic sum, and using the tool box (SOStools) in Matlab into
Row solves, and can effectively overcome dyscalculia.
For this purpose, the present invention devises a kind of control method based on the tiltrotor aircraft mode conversion stage, such as Fig. 1 institute
Show, includes the following steps:
Step 1: establishing longitudinal nonlinear model of tiltrotor aircraft.Its model are as follows:
Wherein, V indicates the actual speed of tiltrotor aircraft,Indicate the one of the actual speed of tiltrotor aircraft
Order derivative, α indicate the actual angle of attack of tiltrotor aircraft,Indicate that the single order of the actual angle of attack of tiltrotor aircraft is led
Number,Indicate the practical pitch angle of tiltrotor aircraft,Indicate that the single order of the practical pitch angle of tiltrotor aircraft is led
Number, q indicate the practical pitch rate of tiltrotor aircraft,Indicate the practical pitch rate of tiltrotor aircraft
First derivative, H indicate the actual height of tiltrotor aircraft,Indicate that the single order of the actual height of tiltrotor aircraft is led
Number, m indicate the quality of tiltrotor aircraft, IzIndicate rotary inertia of the tiltrotor aircraft about pitch axis, FxtIt indicates
Along the practical force component of body x-axis, FytIndicate the practical force component along body y-axis, MzIndicate the practical pitching moment along body,Indicate the state of tiltrotor aircraft, τ (t) indicates that the tilt angle of tiltrotor aircraft, δ indicate tiltrotor aircraft
Longitudinally controlled face.
Step 2: consider the track following problem of tiltrotor aircraft, the longitudinal direction of the tiltrotor aircraft of foundation is non-
Linear model converts error model.
The reference locus of given tiltrotor aircraft, i.e., the speed V of the tiltrotor aircraft of given reference*, the angle of attack
α*, pitch anglePitch rate q*With height H*, and along the power of body x-axisThe power of y-axisAnd pitching momentLongitudinal nonlinear model of the tiltrotor aircraft of foundation is converted into error model:
Wherein, Δ V=V-V*, Δ α=α-α*,Δ q=q-q*, Δ H=H-H*,
Step 3: choosing suitable state variable, control input variable for longitudinal nonlinearity erron of tiltrotor aircraft
Model conversation is the form of easy-to-handle state space.
For the ease of the design of nonlinear time-varying state controller, the present invention is longitudinally non-thread to tiltrotor aircraft first
Property model does following processing.
Choose state variable x1=Δ V, x2=Δ α,x4=Δ q, x5=Δ H, and control input variable u1=Δ
Fxt, u2=Δ Fyt, u3=Δ Mz.In order to facilitate calculating, and assume sin α ≈ α, cos α ≈ 1,With
And nonpolynomialIt is fitted to polynomial form, i.e.,Then it will vert
Rotor craft longitudinal direction nonlinear model is converted into following state space form:
Wherein,
A52=-V*, A53=V*,B11=1/m, B12=-(α*+ Δ α)/m, B43=
1/Iz,
A=8.791 × 10-4, b=-0.03274, c=0.3491.For state vector,To control input vector.
Step 4: by assuming that the speed of tiltrotor aircraft, the angle of attack, pitch angle, pitch rate and height can be surveyed and
Construct nonlinear time-varying state controller.
According to the state space equation (3) of tiltrotor aircraft, the nonlinear time-varying controller of following form is constructed:
Wherein,The gain matrix of controller is controlled for nonlinear time-varying to be designed,Wait ask
Lyapunov matrix, x indicate the state variable chosen,For the group for making state variable that the row vector of corresponding B (x, τ) is 0
It closes, defines J={ j in the present invention1,j2,j3It is the set that the corresponding row matrix of matrix B (x, θ) is 0.τ indicates that tilting rotor flies
The tilt angle of row device has N number of component, is write as τ=[τ1,τ2,…,τN],For tiltrotor aircraft tilt angle to the time
Derivative, it is corresponding to rememberAnd there are the corresponding upper bounds for each componentAnd lower boundvi ；
Step 5: converting the convex optimization of quadratic sum for the design problem of the nonlinear time-varying control controller of tilting rotor and ask
Topic is solved.
Step 5.1: coupled system (3) and nonlinear time-varying controller (4) constitute following closed-loop system:
In the present invention, it would be desirable to closed-loop system (4) locally consistent Exponential Stability at x=0.Therefore, when non-linear
The design problem for becoming controller, which can be converted into, solves unknown gain controlWith Lyapunov matrixMake
Obtain closed-loop system (5) locally consistent Exponential Stability.
Step 5.2: to closed-loop system (5), choosing candidate's Lyapunov function are as follows:
Lyapunov (Liapunov) function against time derivation can obtain:
Wherein, He indicates the sum of number and its transposition inside its bracket.
Step 5.3: by nonlinear time-varying controller (4) and Lyapunov function (6), utilizing Lyapunov stability theorem
And the convex optimum theory of quadratic sum is it is found that in order to reduce conservative, and using S- process, tilting rotor flies-nonlinear time-varying of row device
The design problem of controller can be converted into the following convex optimization inequality of solution:
Wherein,
For Lyapunov matrix to be asked；The gain of controller is controlled for nonlinear time-varying to be asked
Matrix；X indicates state variable,Indicate the state variable so that the row vector all 0 of B (x, τ), τ indicates tilting rotor flight
The tilt angle of device,Indicate derivative of the tilt angle to the time of tiltrotor aircraft；Withvi It respectively indicatesThe upper bound and under
Boundary；I indicates the unit matrix of suitable dimension；z1Indicate suitable dimension non-zero column vector；ε1, ε2And ε3Respectively indicate given normal number.
Step 6: can be obtained using the tool box (SOStools) in Matlab by solving above-mentioned convex optimization problem
One exponentially stable controller of locally consistent (4).
The nonlinear time-varying controller ensure that the locally consistent exponential stability of closed-loop system, and then realize rotation of verting
Control target of the rotor aircraft in the mode conversion stage.
In order to verify the feasibility and validity of method in the present invention, by nonlinear time-varying controller designed by the present invention
It is applied in the control problem in mode conversion stage of certain tiltrotor aircraft.
Simulation parameter: ε1=10-3, ε2=104, ε3=10-3, primary condition: x0=[1 0.01 0.01 0.01 0.1]T,
Simulation result is as shown in Fig. 2-Figure 12, from Fig. 2-Fig. 4 it is known that tilt angle τ,WithIt is all that continuously, tilt angle τ is made
It is meaningful for parameter, and in time t=39s, tiltrotor aircraft completes the conversion from helicopter to aircraft
Mode.From Fig. 5-Fig. 9 it should be apparent that for state variable, although two methods finally all converge to equalization point,
But method proposed by the present invention is capable of fast tracking reference signal, has faster convergence rate.It can be with from Figure 10-Figure 12
Find out, for input variable, method proposed by the present invention has smaller input.Therefore it is proposed in the present invention non-linear
Time-varying control device can be effectively solved tiltrotor aircraft mode conversion problem, effectively complete entire mission phase.
The present invention initially sets up longitudinal nonlinear model of tiltrotor aircraft, according to the track of tiltrotor aircraft
Tracking information, establishes longitudinal Nonlinear Error Models of tiltrotor aircraft, and by choose suitable state variable and
Parameter is converted into longitudinal Nonlinear Error Models of tiltrotor aircraft the form of nonlinear time-varying state space；Its
It is secondary, construct the state controller of nonlinear time-varying；Finally by the design problem of controller be converted into the convex optimization problem of quadratic sum into
Row solves, thus control problem of the effective solution tiltrotor aircraft in the mode conversion stage.
It should be noted that although the above embodiment of the present invention be it is illustrative, this be not be to the present invention
Limitation, therefore the invention is not limited in above-mentioned specific embodiment.Without departing from the principles of the present invention, all
The other embodiment that those skilled in the art obtain under the inspiration of the present invention is accordingly to be regarded as within protection of the invention.
Claims (3)
1. the control method based on the tiltrotor aircraft mode conversion stage, characterized in that specifically include that steps are as follows:
Step 1, the longitudinal nonlinear model for establishing tiltrotor aircraft；
The reference locus of step 2, given tiltrotor aircraft, i.e., the speed of the tiltrotor aircraft of given reference, the angle of attack,
Pitch angle, pitch rate and height, and given reference along the power of body x-axis, the power of y-axis and pitching moment, will establish
Tiltrotor aircraft longitudinal nonlinear model convert error model；
Step 3 chooses state variable x and control input variable u, and longitudinal Nonlinear Error Models of tiltrotor aircraft are turned
Turn to the form of easy-to-handle state space；Wherein
X=[(V-V*)T (α-α*)T (θ-θ*)T (q-q*)T (H-H*)T]T
Step 4, by assuming that the speed of tiltrotor aircraft, the angle of attack, pitch angle, pitch rate and height can be surveyed to construct
Nonlinear time-varying controls controller；
The design problem of the nonlinear time-varying control controller of tiltrotor aircraft is converted the convex optimization of quadratic sum by step 5
Problem is solved, i.e.,
In the above formulas, V and V*Respectively indicate the reality and reference velocity of tiltrotor aircraft；α and α*Respectively indicate rotation of verting
The reality of rotor aircraft and refer to the angle of attack；θ and θ*Respectively indicate the reality of tiltrotor aircraft and with reference to pitch angle；Q and q*Point
Not Biao Shi tiltrotor aircraft reality and refer to pitch rate；H and H*Respectively indicate the reality of tiltrotor aircraft
And reference altitude；FxtWithIt respectively indicates along the reality of body x-axis and with reference to force component；FytWithIt respectively indicates along body y
The reality of axis and refer to force component；MzWithIt respectively indicates along the reality of body and with reference to pitching moment；
A51=Δ θ-Δ α, A52=-V*, A53=V*；Aj(x, τ) expression A (x,
Jth row τ), j=1,2 ..., 5；
B11=1/m, B12=-(α*+ Δ α)/m, B43=1/Iz,
A=8.791 × 10-4, b=-0.03274, c=0.3491；
For Lyapunov matrix to be asked；The gain square of controller is controlled for nonlinear time-varying to be asked
Battle array；X indicates state variable,Indicate the state variable so that the row vector all 0 of B (x, τ), τ indicates tiltrotor aircraft
Tilt angle,Indicate derivative of the tilt angle to the time of tiltrotor aircraft；Withv iIt respectively indicatesThe upper bound and lower bound；
I indicates the unit matrix of suitable dimension；z1Indicate suitable dimension non-zero column vector；ε1, ε2And ε3Respectively indicate given normal number；
Step 6, using solve come nonlinear time-varying control controller tiltrotor aircraft mode conversion can be realized
The control target in stage.
2. the control method according to claim 1 based on the tiltrotor aircraft mode conversion stage, characterized in that step
In 4, the nonlinear time-varying state controller that is constructed.
3. the control method according to claim 1 based on the tiltrotor aircraft mode conversion stage, characterized in that step
In 5, the convex optimization problem of quadratic sum is solved using the tool box SOStools in Matlab.
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Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104460681A (en) * | 2014-09-24 | 2015-03-25 | 南京航空航天大学 | Method for controlling flying of tilt-rotor unmanned helicopter in transition section |
| CN108052008A (en) * | 2018-01-10 | 2018-05-18 | 大连理工大学 | A kind of geometry controller design method of tiltrotor aircraft transition state handoff procedure |
| CN108427272A (en) * | 2018-03-13 | 2018-08-21 | 厦门大学 | Flexible Spacecraft control based on flexible mode observation and vibration suppressing method |
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2019
- 2019-01-23 CN CN201910062363.0A patent/CN109581878B/en not_active Expired - Fee Related
Patent Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104460681A (en) * | 2014-09-24 | 2015-03-25 | 南京航空航天大学 | Method for controlling flying of tilt-rotor unmanned helicopter in transition section |
| CN108052008A (en) * | 2018-01-10 | 2018-05-18 | 大连理工大学 | A kind of geometry controller design method of tiltrotor aircraft transition state handoff procedure |
| CN108427272A (en) * | 2018-03-13 | 2018-08-21 | 厦门大学 | Flexible Spacecraft control based on flexible mode observation and vibration suppressing method |
Non-Patent Citations (3)
| Title |
|---|
| RONG FU等: "Exponential stabilisation of nonlinear parameter-varying systems with applications to conversion flight control of a tilt rotor aircraft", 《INTERNATIONAL JOURNAL OF CONTROL》 * |
| STEPHEN PRAJNA等: "《Nonlinear Control Synthesis by Sum of Squares》", 《2004 5TH ASIAN CONTROL CONFERENCE》 * |
| 鲁麟宏等: "基于矫正广义走廊的电动倾转旋翼机模态转换", 《航空学报》 * |
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