CN104460681A - Method for controlling flying of tilt-rotor unmanned helicopter in transition section - Google Patents

Abstract

A method for controlling flying of a tilt-rotor unmanned helicopter in a transition section is characterized in that the optimal preview control is adopted, and by fusing system motion equation constraints, a preview ideal state and the optimal performance index constraint information, the flying of the tilt-rotor unmanned helicopter in the transition section is controlled; specifically, the method comprises the steps that firstly, a nonlinear model is set up for the transition section of the tilt-rotor unmanned helicopter, and then a liner equation for state space description is obtained in the mode that trimming and linearization at different nacelle dip angles are achieved; secondly, based on the linear model, the optimal preview controller is designed, and the flying speed, the flying height and the flying attitude angle in the transition process are controlled; finally, inertia delay desalted equipment is designed for achieving smooth transition of the tilt-rotor unmanned helicopter between different nacelle dip angles. According to the method for controlling flying of the tilt-rotor unmanned helicopter in the transition section, by fusing a predicable expected speed, a height, an attitude angle and other information, safe switching between modes of the tilt-rotor unmanned helicopter is achieved.

Description

The flight control method of tilting rotor depopulated helicopter transition section

Technical field

The present invention relates to the flight control method of tilting rotor depopulated helicopter transition section, particularly relate to a kind of flight control method of Optimum-preview control, belong to technical field of flight control.

Background technology

Tiltrotor is a kind of mixing aircraft, both can as fixed wing aircraft before fly, while also can realize vertical takeoff and landing and hovering as helicopter.The property of tiltrotor, extends it in dual-use middle usable range.The U.S. " Eagle Eye " is a tilting rotor depopulated helicopter developed at present.

Tiltrotor brings many new problems and technological difficulties: on the one hand, aircraft control mode and offline mode more complicated, the model adding transition section describes difficulty, existing airvane controls there is again pulling force vector controlled, not only there are helicopter mode, fixed-wing pattern, there is again the handoff procedure between helicopter mode and fixed-wing pattern.On the other hand, in mode handover procedure, Aerodynamic Characteristics is extremely complicated, and has stronger pitching couple and pitching/plunging motion coupling when engine nacelle verts, and Aerodynamic intenference is serious, and aircraft stability is poor.Therefore, the switching how making aircraft stably complete between tiltrotor pattern is even more important for safe flight.And the requirement of tilting rotor depopulated helicopter to flight control system to have compared people's tiltrotor more harsh.

For the research of tiltrotor transition section, the particularly research of transition section control program, Chinese scholars proposes multiple control method.Except the classical control method of majority, Dr.Anthony J.Calise and Rolf T.Rysdyk proposes adaptive nonlinear control method, and the method adopts network and models coupling.But these control methods often only make use of the past and present information of system.But control for aircraft, Future Trajectory information realizes planning and design.Foresee controlling is not only considered the information that system is current also to decide current control program according to the Future Information of system, therefore can improve the control performance of system.

At present, foresee controlling successful Application in some fields.As mechanical hook-up initiatively damping time, can apply foresee controlling before vibration produces or undesired signal occur before with regard to access control effect, make control energy reduce a lot; And for example in the path following control of lathe, robot etc., because path is known mostly in advance, application foresee controlling can greatly reduce track following error.And it is little in the application of flying vehicles control technical field.

Summary of the invention

Technical matters to be solved:

The object of this invention is to provide a kind of flight control method of tilting rotor depopulated helicopter transition section, merge the information such as foreseeable desired trajectory and desired speed, complete the handoff-security between tilting rotor depopulated helicopter mode.

Technical scheme:

In order to realize above function, the invention provides a kind of flight control method of tilting rotor depopulated helicopter transition section, it is characterized in that, the method adopts Optimum-preview control by the constraint of the emerging system equation of motion, foreseeable perfect condition and optimal performance index constraint information, controls the flight of tilting rotor depopulated helicopter transition section; Concrete grammar is:

First set up nonlinear model to tilting rotor depopulated helicopter transition section, then under different nacelle inclination angle, trim linearization obtains the linear equation of state space description;

Secondly, based on linear model, devise optimum Preview Controller, controls flying speed, height and the attitude angle in transient process;

Finally seamlessly transit Inertia of design delayed type desalination apparatus for tilting rotor depopulated helicopter is formed between different nacelle inclination angle.

Concretely, described different nacelle angled state is β _m0 °: 15 °, 15 °: 30 °, 30 °: 45 °, 45 °: 60 °, 60 °: 75 °, 75 °: 90 °; At β _mbe 15 °, 30 °, 45 °, 60 °, 75 °, 90 °, set up the linear model of the tilting rotor depopulated helicopter transition section corresponding to above-mentioned six stages, be expressed as A1, B1, C1, D1}, A2, B2, C2, D2} ..., { A6, B6, C6, D6}}; Wherein, above-mentioned Ai, Bi, Ci, Di represent the system matrix of linear model respectively, described every one-phase { Ai, Bi, Ci, Di} comprise longitudinal model { Ai _lon, Bi _lon, Ci _lon, Di _lonand horizontal side direction model { Ai _lat, Bi _lat, Ci _lat, Di _lat.

Preferred as one, when described nacelle inclination angle is no more than 45 °, longitudinal mode of tilting rotor depopulated helicopter transition section is total apart from δ _cwith longitudinal feathering δ _lonjointly control, horizontal side direction is horizontal feathering δ _latwith longitudinal differential feathering δ _dlonjointly control; Then front three phaseslongitudinal Control Law form be

$\mathrm{\ΔU}\left(k\right)=\left[\begin{array}{c}{\mathrm{\δ}}_c\\ {\mathrm{\δ}}_{\mathrm{lon}}\end{array}\right]=F_0{X}_0\left(k\right)+\underset{j=0}{\overset{M_R}{\mathrm{\Σ}}}{F}_R\left(j\right)\mathrm{\ΔR}(k+j)$

X ₀(k)＝[*u *w *q u w q θ H] ^T，R(k)＝[u w q] ^T

In formula, u is the flying speed along body coordinate x-axis direction, and w is the flying speed along body coordinate z-axis direction, and q is rate of pitch, adds the difference that * represents this variable measurements and ideal value before variable; δ _cfor total distance, δ _lonfor longitudinal feathering, θ is the angle of pitch; H is flying height;

front three phaseshorizontal side direction control law form be

$\mathrm{\ΔU}\left(k\right)=\left[\begin{array}{c}{\mathrm{\δ}}_{\mathrm{lat}}\\ {\mathrm{\δ}}_{\mathrm{dlon}}\end{array}\right]=F_0{X}_0\left(k\right)+\underset{j=0}{\overset{M_R}{\mathrm{\Σ}}}{F}_R\left(j\right)\mathrm{\ΔR}(k+j)$

X ₀(k)＝[*β *p *r β p r φ] ^T，R(k)＝[β p r] ^T。

In formula, β is yaw angle, and p is angular velocity in roll, and r is yaw rate, and φ is roll angle, adds the difference that * represents this variable measurements and ideal value before variable; δ _latfor horizontal feathering, δ _dlonfor longitudinal differential feathering

Preferred as one, after described nacelle inclination angle is more than 45 °, longitudinal mode of tilting rotor depopulated helicopter transition section is total apart from δ _cwith elevating rudder δ _ejointly control, horizontal side direction is aileron δ _awith yaw rudder δ _rjointly control; Then, the Longitudinal Control Law form of three phases is

$\mathrm{\ΔU}\left(k\right)=\left[\begin{array}{c}{\mathrm{\δ}}_c\\ {\mathrm{\δ}}_e\end{array}\right]=F_0{X}_0\left(k\right)+\underset{j=0}{\overset{M_R}{\mathrm{\Σ}}}{F}_R\left(j\right)\mathrm{\ΔR}(k+j)$

X ₀(k)＝[*u *w *q u w q θ H] ^T，R(k)＝[u w q] ^T

In formula, δ _efor elevating rudder drift angle;

The horizontal side direction control law form of rear three phases is

$\mathrm{\ΔU}\left(k\right)=\left[\begin{array}{c}{\mathrm{\δ}}_a\\ {\mathrm{\δ}}_r\end{array}\right]=F_0{X}_0\left(k\right)+\underset{j=0}{\overset{M_R}{\mathrm{\Σ}}}{F}_R\left(j\right)\mathrm{\ΔR}(k+j)$

X ₀(k)＝[*β *p *r β p r φ] ^T，R(k)＝[β p r] ^T。

In formula, δ _afor aileron drift angle, δ _rfor rudder.

Concretely, the method for designing of inertial delay formula desalination apparatus is specially:

If the output of former control law is U _at (), the output of reconfigurable control rule is U _b(t), the output of switching law is U (t); Suppose at t ₀moment carries out control laws transformation, to t ₁moment, through exponential damping effect, and complete control laws transformation.The specific algorithm of this type desalination apparatus is

$U\left(t\right)=\left\{\begin{array}{cc}{U}_A\left(t\right)& t\&le;t_0\\ U_A\left(t\right)e^{-a(t-t_0)}+U_B\left(t\right)(1-e^{-a\left({t-t}_0\right)})& t_0<t\&le;t_1\\ U_B\left(t\right)& t>t_1\end{array}\right.$

Parameter a is the principal element of impact desalination performance, and its value is larger, and fade time is shorter, otherwise its value is less, and fade time is longer.

Beneficial effect:

The object of this invention is to provide a kind of flight control method of tilting rotor depopulated helicopter transition section, merge foreseeable desired speed, height and attitude angle information, complete the handoff-security between tilting rotor depopulated helicopter mode.

The flight control method of a kind of tilting rotor depopulated helicopter transition section of the present invention, adopt Optimum-preview control by information such as emerging system equation of motion hard constraint, foreseeable perfect condition soft-constraint and optimal performance index soft-constraints, the flight of tilting rotor depopulated helicopter transition section is controlled.

Accompanying drawing explanation

Below in conjunction with accompanying drawingthe present invention is further described with embodiment:

fig. 1for tilting rotor depopulated helicopter transition section speed safe range;

fig. 2for tilting rotor depopulated helicopter transition section of the present invention front three phaseslongitudinal Optimum-preview control device structure;

fig. 3for control laws transformation desalination apparatus simulation curve of the present invention figure;

fig. 4the longitudinal velocity response curve at 15 degree, the nacelle inclination angle obtained for adopting flight control method of the present invention;

fig. 5the longitudinal velocity response curve at 45 degree, the nacelle inclination angle obtained for adopting flight control method of the present invention;

fig. 6the longitudinal velocity response curve at 60 degree, the nacelle inclination angle obtained for adopting flight control method of the present invention;

fig. 7the one dimension track emulation curve obtained for adopting flight control method of the present invention;

fig. 8the two-dimentional track emulation curve obtained for adopting flight control method of the present invention.

Embodiment

The invention provides a kind of optimum prediction flight control method of tilting rotor depopulated helicopter transition section, for making object of the present invention, technical scheme and effect clearly, clearly, and reference accompanying drawingthe present invention is described in more detail for examples.Should be appreciated that concrete enforcement described herein is only in order to explain the present invention, is not intended to limit the present invention.

Adopt the optimum prediction flight control method of tilting rotor depopulated helicopter transition section of the present invention, specifically according to following steps.

Step 1, first set up the nonlinear model of tilting rotor depopulated helicopter transition section.Under body axis system, vehicle dynamics and kinematics six degree of freedom system of equations can adopt following form.

$X=m(\stackrel{\&CenterDot;}{u}+\mathrm{qw}-\mathrm{rv})+\mathrm{mg}\sin \mathrm{\&theta;}$

$Y=m(\stackrel{\&CenterDot;}{v}+\mathrm{ru}-\mathrm{pw})-\mathrm{mg}\sin \mathrm{\&phi;}\cos \mathrm{\&theta;}$

$Z=m(\stackrel{\&CenterDot;}{w}+\mathrm{pv}-\mathrm{qu})-\mathrm{mg}\cos \mathrm{\&phi;}\cos \mathrm{\&theta;}$

$L=I_{\mathrm{xx}}\stackrel{\&CenterDot;}{p}-(I_{\mathrm{yy}}-I_{\mathrm{zz}})\mathrm{qr}+I_{\mathrm{yz}}(r^2-q^2)-I_{\mathrm{xz}}(\mathrm{pq}+\stackrel{\&CenterDot;}{r})+I_{\mathrm{xy}}(\mathrm{pr}-\stackrel{\&CenterDot;}{q})$

$M=I_{\mathrm{yy}}\stackrel{\&CenterDot;}{q}-(I_{\mathrm{zz}}-I_{\mathrm{xx}})\mathrm{pr}+I_{\mathrm{xz}}(p^2-r^2)-I_{\mathrm{xy}}(\mathrm{qr}+\stackrel{\&CenterDot;}{p}){+I}_{\mathrm{yz}}(\mathrm{pq}-r^2)$

$N=I_{\mathrm{zz}}\stackrel{\&CenterDot;}{r}-(I_{\mathrm{xx}}-I_{\mathrm{yy}})\mathrm{pq}+I_{\mathrm{xy}}(q^2-p^2)-I_{\mathrm{yz}}(\mathrm{pr}+\stackrel{\&CenterDot;}{q})+I_{\mathrm{xz}}(\mathrm{qr}-\stackrel{\&CenterDot;}{p})$

$p=\stackrel{\&CenterDot;}{\mathrm{\&phi;}}-\stackrel{\&CenterDot;}{\mathrm{\&psi;}}\cos \mathrm{\&theta;}$

$q=\stackrel{\&CenterDot;}{\mathrm{\&theta;}}\cos \mathrm{\&phi;}+\stackrel{\&CenterDot;}{\mathrm{\&psi;}}\sin \mathrm{\&phi;}\cos \mathrm{\&theta;}$

$r=\stackrel{\&CenterDot;}{\mathrm{\&psi;}}\cos \mathrm{\&phi;}\cos \mathrm{\&theta;}-\stackrel{\&CenterDot;}{\mathrm{\&theta;}}\sin \mathrm{\&phi;}$

When calculating force and moment, the effect of every part such as fuselage, rotor, wing, horizontal tail and vertical fin be considered.As being constructed as follows of side force in axis y-axis direction

Y＝Y _fuse+Y _wing+Y _ROT+Y _HT+Y _VT

In formula, Y _fusethe side force that fuselage produces, Y _wingthe side force that wing produces, Y _rOTthe side force that rotor produces, Y _hTthe side force that horizontal tail produces, Y _vTit is the side force that vertical fin produces.

After respectively mathematical modeling being carried out to fuselage, rotor, wing, horizontal tail and vertical fin, in MATLAB/Simulink, build the nonlinear mathematical model of tilting rotor depopulated helicopter.

Then the trim that the trim function in MATLAB/Simulink realizes to system under different nacelle angle is used, recycling linmod function, using trim input quantity and the trim condition amount parameter as linmod function, by nonlinear model linearization near trim point.Because the motion of tilting rotor depopulated helicopter transition section has very strong time-varying characteristics, transition section is divided six stages, respectively corresponding nacelle angle of inclination beta _m0 °: 15 °, 15 °: 30 °, 30 °: 45 °, 45 °: 60 °, 60 °: 75 °, 75 °: 90 °.At β _mbe 15 °, 30 °, 45 °, 60 °, 75 °, 90 ° of foundation correspond to the linear model of the tilting rotor depopulated helicopter transition section in above-mentioned six stages, be expressed as A1, B1, C1, D1}, A2, B2, C2, D2} ..., { A6, B6, C6, D6}}.Wherein every one-phase { Ai, Bi, Ci, Di} comprise with x=[u w q θ H] ^tfor state variable, with U=[δ _cδ _lonδ _e] be longitudinal model { Ai of controlled quentity controlled variable _lon, Bi _lon, Ci _lon, Di _lonand with x=[β p r φ] ^tfor state variable, with U=[δ _latδ _dlonδ _aδ _r] be the horizontal side direction model { Ai of controlled quentity controlled variable _lat, Bi _lat, Ci _lat, Di _lat.

Understand technical scheme of the present invention for the ease of the public, use following table 1the physical quantity that the present invention relates to is described:

Variable	Implication
		u	Along the flying speed in body coordinate x-axis direction
w	Along the flying speed in body coordinate z-axis direction
		β M	Nacelle inclination angle
θ	The angle of pitch
		φ	Roll angle
ψ	Crab angle
		p	Angular velocity in roll
q	Rate of pitch
		r	Yaw rate
δ c	Total distance
		δ lon	Longitudinal feathering

δ dlon	Longitudinal differential feathering
		δ lat	Horizontal feathering
δ e	Elevating rudder drift angle
		δ a	Aileron drift angle
δ r	Rudder
		m	Vehicle mass
β	Yaw angle
		φ	Roll angle
p	Angular velocity in roll
		r	Yaw rate
δ a	Aileron drift angle
		δ r	Rudder
I xx	Aircraft is around the moment of inertia of body axis system x-axis
		I yy	Aircraft is around the moment of inertia of body axis system y-axis
I zz	Aircraft is around the moment of inertia of body axis system z-axis
		I xy	Aircraft is to the product of inertia of body axis system x, y-axis
I xz	Aircraft is to the product of inertia of body axis system x, z-axis
		I yz	Aircraft is to the product of inertia of body axis system y, z-axis
X	The making a concerted effort along body axis system x-axis that aircraft is subject to
		Y	The side force along body axis system y-axis that aircraft is subject to
Z	The making a concerted effort along body axis system z-axis that aircraft is subject to
		L	Rolling moment
M	Pitching moment
		N	Yawing
Δ is added before variable	The increment of this variable
		* is added before variable	The difference of this variable measurements and ideal value

table 1

Step 2, based on above-mentioned six linear models, devise optimum Preview Controller.

First control program is determined, as follows:

(1) when nacelle inclination angle is no more than 45 °, longitudinal mode of tilting rotor depopulated helicopter transition section is total apart from δ _cwith longitudinal feathering δ _lonjointly control, horizontal side direction is horizontal feathering δ _latwith longitudinal differential feathering δ _dlonjointly control.

(2) nacelle inclination angle more than 45 ° after, longitudinal mode of tilting rotor depopulated helicopter transition section is total apart from δ _cwith elevating rudder δ _ejointly control, horizontal side direction is aileron δ _awith yaw rudder δ _rjointly control.

According to above-mentioned control program, the state equation in the stage of six in step 1 is divided, obtains front three phaseslengthwise movement state equation be

$\left\{\begin{array}{c}{\left[\begin{array}{ccccc}\stackrel{\&CenterDot;}{u}& \stackrel{\&CenterDot;}{w}& \stackrel{\&CenterDot;}{q}& \stackrel{\&CenterDot;}{\mathrm{\&theta;}}& \stackrel{\&CenterDot;}{H}\end{array}\right]}^T=A_{\mathrm{lon}}^{\&prime;}{\left[\begin{array}{ccccc}u& w& q& \mathrm{\&theta;}& H\end{array}\right]}^T+B_{\mathrm{lon}}^{\&prime;}{\left[\begin{array}{cc}{\mathrm{\&delta;}}_c& {\mathrm{\&delta;}}_{\mathrm{lon}}\end{array}\right]}^T\\ {\left[\begin{array}{ccccc}u& w& q& \mathrm{\&theta;}& H\end{array}\right]}^T{=C}_{\mathrm{lon}}^{\&prime;}{\left[\begin{array}{ccccc}u& w& q& \mathrm{\&theta;}& H\end{array}\right]}^T\end{array}\right.$

front three phaseshorizontal sideway movement state equation be

$\left\{\begin{array}{c}{\left[\begin{array}{cccc}\stackrel{\&CenterDot;}{\mathrm{\&beta;}}& \stackrel{\&CenterDot;}{p}& \stackrel{\&CenterDot;}{r}& \stackrel{\&CenterDot;}{\mathrm{\&phi;}}\end{array}\right]}^T=A_{\mathrm{lat}}^{\&prime;}{\left[\begin{array}{cccc}\mathrm{\&beta;}& p& r& \mathrm{\&phi;}\end{array}\right]}^T{+B}_{\mathrm{lat}}^{\&prime;}{\left[\begin{array}{cc}{\mathrm{\&delta;}}_{\mathrm{lat}}& {\mathrm{\&delta;}}_{\mathrm{dlon}}\end{array}\right]}^T\\ {\left[\begin{array}{cccc}\mathrm{\&beta;}& p& r& \mathrm{\&phi;}\end{array}\right]}^T=C_{\mathrm{lat}}^{\&prime;}{\left[\begin{array}{cccc}\mathrm{\&beta;}& p& r& \mathrm{\&phi;}\end{array}\right]}^T\end{array}\right.$

The lengthwise movement state equation of rear three phases is

$\left\{\begin{array}{c}{\left[\begin{array}{ccccc}\stackrel{\&CenterDot;}{u}& \stackrel{\&CenterDot;}{w}& \stackrel{\&CenterDot;}{q}& \stackrel{\&CenterDot;}{\mathrm{\&theta;}}& \stackrel{\&CenterDot;}{H}\end{array}\right]}^T=A_{\mathrm{lon}}^{\&prime;\&prime;}{\left[\begin{array}{ccccc}u& w& q& \mathrm{\&theta;}& H\end{array}\right]}^T+B_{\mathrm{lon}}^{\&prime;\&prime;}{\left[\begin{array}{cc}{\mathrm{\&delta;}}_c& {\mathrm{\&delta;}}_{\mathrm{lon}}\end{array}\right]}^T\\ {\left[\begin{array}{ccccc}u& w& q& \mathrm{\&theta;}& H\end{array}\right]}^T{=C}_{\mathrm{lon}}^{\&prime;\&prime;}{\left[\begin{array}{ccccc}u& w& q& \mathrm{\&theta;}& H\end{array}\right]}^T\end{array}\right.$

The horizontal sideway movement state equation of rear three phases is

$\left\{\begin{array}{c}{\left[\begin{array}{cccc}\stackrel{\&CenterDot;}{\mathrm{\&beta;}}& \stackrel{\&CenterDot;}{p}& \stackrel{\&CenterDot;}{r}& \stackrel{\&CenterDot;}{\mathrm{\&phi;}}\end{array}\right]}^T=A_{\mathrm{lat}}^{\&prime;\&prime;}{\left[\begin{array}{cccc}\mathrm{\&beta;}& p& r& \mathrm{\&phi;}\end{array}\right]}^T{+B}_{\mathrm{lat}}^{\&prime;\&prime;}{\left[\begin{array}{cc}{\mathrm{\&delta;}}_{\mathrm{lat}}& {\mathrm{\&delta;}}_{\mathrm{dlon}}\end{array}\right]}^T\\ {\left[\begin{array}{cccc}\mathrm{\&beta;}& p& r& \mathrm{\&phi;}\end{array}\right]}^T=C_{\mathrm{lat}}^{\&prime;\&prime;}{\left[\begin{array}{cccc}\mathrm{\&beta;}& p& r& \mathrm{\&phi;}\end{array}\right]}^T\end{array}\right.$

With certain sampling time, above-mentioned linear model is carried out discretize.Obtain the linear discrete time system that following formula represents

x(k+1)＝Ax(k)+BU(k)

y(k)＝Cx(k)

If target value signal is R (k), definition error signal is

e(k)＝R(k)-y(k)

Then error signal system is below derived

$\left[\begin{array}{c}e(k+1)\\ \mathrm{\&Delta;x}(k+1)\end{array}\right]=\left[\begin{array}{cc}{I}_m& -\mathrm{CA}\\ 0& A\end{array}\right]\left[\begin{array}{c}e\left(k\right)\\ \mathrm{\&Delta;x}\left(k\right)\end{array}\right]+\left[\begin{array}{c}-\mathrm{CB}\\ B\end{array}\right]\mathrm{\&Delta;U}\left(k\right)+\left[\begin{array}{c}{I}_m\\ 0\end{array}\right]\mathrm{\&Delta;R}(k+1)$

Or

X ₀(k+1)＝ΦX ₀(k)+GΔU(k)+G _RΔR(k+1)

For error system, suppose known its from present moment until following M _rthe desired value of step, definition evaluation function is the quadratic form comprising error term and input item as follows:

$\begin{array}{c}J=\underset{k=-M_R+1}{\overset{\&infin;}{\mathrm{\&Sigma;}}}\left[e^T\left(k\right)Q_{e}e\left(k\right){+\mathrm{\&Delta;U}}^T\left(k\right)\mathrm{H\&Delta;U}\left(k\right)\right]\\ =\underset{k=-M_R+1}{\overset{\&infin;}{\mathrm{\&Sigma;}}}[{X_0}^T\left(k\right)Q{X}_0\left(k\right)+{\mathrm{\&Delta;U}}^T\left(k\right)\mathrm{H\&Delta;U}\left(k\right)]\end{array}$

Wherein: Q is positive semidefinite matrix; Q _e, H is positive definite matrix.Q and H is the weight matrix determined by deviser.

Several solution is had for the optimal control problem solving error system, conventional one is that the method is the form first supposing control inputs according to partial differential optimization method, substitutes in evaluation function, require that this function gets minimum value according to partial differential, thus obtain prediction feed-forward coefficients.

If the control inputs form of Optimum-preview control is

$\mathrm{\&Delta;U}\left(k\right)=F_0{X}_0\left(k\right)+\underset{j=0}{\overset{M_R}{\mathrm{\&Sigma;}}}{F}_R\left(j\right)\mathrm{\&Delta;R}(k+j)$

Here, F _rj () is coefficient undetermined.Said method solving result is utilized to be

F ₀＝-[H+G ^TPG] ^-1G ^TPΦ

F _R(j)＝-[H+G ^TPG] ^-1G ^T(ξ ^T) ^j-1PG _R(j≥1)

(ξ＝Φ+GF ₀)

P＝Q+Φ ^TPΦ-Φ ^TPG[H+G ^TPG] ^-1G ^TPΦ

If make F ₀=[F _e, F _x], system initial value is zero, obtains

$U\left(k\right)=F_e\frac{z}{z-1}e\left(k\right)+F_{x}x\left(k\right)+U(k-1)+\underset{j=1}{\overset{M_R}{\mathrm{\&Sigma;}}}{F}_R\left(j\right)R(k+j)+\underset{j=0}{\overset{M_R}{\mathrm{\&Sigma;}}}{F}_d\left(j\right)d(k+j)$

Be to make aircraft forward flight speed within this stage reach the scope of safe transition in the object of each phase controller, be namely in as Fig. 1in shown safe range.Controller architecture in conjunction with controlling plan design is

front three phaseslongitudinal Control Law form be

$\mathrm{\&Delta;U}\left(k\right)=\left[\begin{array}{c}{\mathrm{\&delta;}}_c\\ {\mathrm{\&delta;}}_{\mathrm{lon}}\end{array}\right]=F_0{X}_0\left(k\right)+\underset{j=0}{\overset{M_R}{\mathrm{\&Sigma;}}}{F}_R\left(j\right)\mathrm{\&Delta;R}(k+j)$

X ₀(k)＝[*u *w *q u w q θ H] ^T，R(k)＝[u w q] ^T

front three phaseslongitudinal controller structure figure is as Fig. 2shown in.

front three phaseshorizontal side direction control law form be

$\mathrm{\&Delta;U}\left(k\right)=\left[\begin{array}{c}{\mathrm{\&delta;}}_{\mathrm{lat}}\\ {\mathrm{\&delta;}}_{\mathrm{dlon}}\end{array}\right]=F_0{X}_0\left(k\right)+\underset{j=0}{\overset{M_R}{\mathrm{\&Sigma;}}}{F}_R\left(j\right)\mathrm{\&Delta;R}(k+j)$

X ₀(k)＝[*β *p *r β p r φ] ^T，R(k)＝[β p r] ^T

The Longitudinal Control Law form of rear three phases is

$\mathrm{\&Delta;U}\left(k\right)=\left[\begin{array}{c}{\mathrm{\&delta;}}_c\\ {\mathrm{\&delta;}}_e\end{array}\right]=F_0{X}_0\left(k\right)+\underset{j=0}{\overset{M_R}{\mathrm{\&Sigma;}}}{F}_R\left(j\right)\mathrm{\&Delta;R}(k+j)$

X ₀(k)＝[*u *w *q u w q θ H] ^T，R(k)＝[u w q] ^T

The horizontal side direction control law form of rear three phases is

$\mathrm{\&Delta;U}\left(k\right)=\left[\begin{array}{c}{\mathrm{\&delta;}}_a\\ {\mathrm{\&delta;}}_r\end{array}\right]=F_0{X}_0\left(k\right)+\underset{j=0}{\overset{M_R}{\mathrm{\&Sigma;}}}{F}_R\left(j\right)\mathrm{\&Delta;R}(k+j)$

X ₀(k)＝[*β *p *r β p r φ] ^T，R(k)＝[β p r] ^T

Realize seamlessly transitting between step 3, six phase controller, devise inertial delay formula desalination apparatus for this reason, if the output of former control law is U _at (), the output of reconfigurable control rule is U _bt (), the output of switching law is U (t).Suppose at t ₀moment carries out control laws transformation, to t ₁moment, through exponential damping effect, and complete control laws transformation.The specific algorithm of this type desalination apparatus is

$U\left(t\right)=\left\{\begin{array}{cc}{U}_A\left(t\right)& t\&le;t_0\\ U_A\left(t\right)e^{-a(t-t_0)}+U_B\left(t\right)(1-e^{-a\left({t-t}_0\right)})& t_0<t\&le;t_1\\ U_B\left(t\right)& t>t_1\end{array}\right.$

Parameter a is the principal element of impact desalination performance, and its value is larger, and fade time is shorter, otherwise its value is less, and fade time is longer.

Suppose that switching instant is 5s, for the controlled quentity controlled variable of input, be-5 before switching, be 8 after switching, a=1.5 here, switching time is 2s, then the handoff procedure of desalination apparatus as Fig. 3shown in.Now the algorithm of desalination apparatus is as follows

$u\left(t\right)=\left\{\begin{array}{cc}-5& t\&le;5\\ {-5e}^{-1.5(t-5)}+8\left({1-e}^{-1.5(t-5)}\right)& 5<t\&le;7\\ 8& t>7\end{array}\right.$

In order to verify the validity that the present invention controls in the flight of tilting rotor transition section, carry out following emulation experiment.Emulation tool adopts MATLAB software, and object adopts XV-15 tiltrotor aircraft to analyze.Simulation parameter is arranged as table 2shown in.

Time t (s)	0	0～5	5～15	15～55
					β M	0°	15°	30°	60°

u(m/s)	20	30	50	60
					w(m/s)	2.85	3.45	4.65	7.6

table 2

With β in 0 ~ 5s _mtrim linear model when=15 ° controls, with β in 5 ~ 15s _mtrim linear model when=30 ° controls, with β in 15 ~ 55s _mtrim linear model when=60 ° controls, table 2the setting of middle forward flight speed u drops on safely fig. 1in shown safe range.The tracking mode arranging speed is in front 15s, and u is with acceleration a _u=2m/s ², w is with acceleration a _w=0.12m/s ²climb; Inputting amplitude at 15s moment u is the step signal of 10, and it is the step signal of 2.95 that w inputs amplitude.If elemental height is 100m, according to above-mentioned steps design con-trol Optimum-preview control device, control effects as Fig. 4arrive fig. 8shown in, wherein fig. 7, all stage track shown in 8 according to fig. 4arrive fig. 6rate curve calculate.? fig. 4arrive fig. 6ordinate represent that present speed deducts the difference of initial velocity, fig. 4in, although steady-state error appears in the control result of u, within the scope of safe transition, the transition of tilting rotor can not be affected.? fig. 6the control of middle w to have the process declined time initial, but is greater than 0 due to w initial value, so there will be the height decline phenomenon in the short time, can recover climb mode afterwards, from fig. 7altitude curve change in also can find out.The object controlled in this emulation is mainly speeds control, is not track following, and therefore track can have deviation with ideal trajectory.Above-mentioned simulation result fully shows, has good response performance based on there being most the tilting rotor unmanned helicopter system of foresee controlling.

Be understandable that, for those of ordinary skills, can be equal to according to technical scheme of the present invention and inventive concept thereof and replace or change, and all these change or replace the protection domain that all should belong to the claim appended by the present invention.

Claims (5)

1. one kind the flight control method of tilting rotor depopulated helicopter transition section, it is characterized in that, the method adopts Optimum-preview control by the constraint of the emerging system equation of motion, foreseeable perfect condition and optimal performance index constraint information, controls the flight of tilting rotor depopulated helicopter transition section; Concrete grammar is:

First set up nonlinear model to tilting rotor depopulated helicopter transition section, then under different nacelle inclination angle, trim linearization obtains the linear equation of state space description;

Secondly, based on linear model, devise optimum Preview Controller, controls flying speed, height and the attitude angle in transient process;

Finally seamlessly transit Inertia of design delayed type desalination apparatus for tilting rotor depopulated helicopter is formed between different nacelle inclination angle.

2. one according to claim 1 the flight control method of tilting rotor depopulated helicopter transition section, it is characterized in that: described different nacelle angled state is β _m0 °: 15 °, 15 °: 30 °, 30 °: 45 °, 45 °: 60 °, 60 °: 75 °, 75 °: 90 °; At β _mbe 15 °, 30 °, 45 °, 60 °, 75 °, 90 °, set up the linear model of the tilting rotor depopulated helicopter transition section corresponding to above-mentioned six stages, be expressed as A1, B1, C1, D1}, A2, B2, C2, D2} ..., { A6, B6, C6, D6}}; Wherein, above-mentioned Ai, Bi, Ci, Di represent the system matrix of linear model respectively, described every one-phase { Ai, Bi, Ci, Di} comprise longitudinal model { Ai _lon, Bi _lon, Ci _lon, Di _lonand horizontal side direction model { Ai _lat, Bi _lat, Ci _lat, Di _lat.

3. the flight control method of a kind of tilting rotor depopulated helicopter transition section according to claim 2, is characterized in that: when described nacelle inclination angle is no more than 45 °, and longitudinal mode of tilting rotor depopulated helicopter transition section is total apart from δ _cwith longitudinal feathering δ _lonjointly control, horizontal side direction is horizontal feathering δ _latwith longitudinal differential feathering δ _dlonjointly control; Then the Longitudinal Control Law form of front three phases is

X ₀(k)＝[*u *w *q u w q θ H] ^T，R(k)＝[u w q] ^T

In formula, u is the flying speed along body coordinate x-axis direction, and w is the flying speed along body coordinate z-axis direction, and q is rate of pitch, adds the difference that * represents this variable measurements and ideal value before variable; δ _cfor total distance, δ _lonfor longitudinal feathering, θ is the angle of pitch; H is flying height;

The horizontal side direction control law form of front three phases is

X ₀(k)＝[*β *p *r β p r φ] ^T，R(k)＝[β p r] ^T。

In formula, β is yaw angle, and p is angular velocity in roll, and r is yaw rate, and φ is roll angle, adds the difference that * represents this variable measurements and ideal value before variable; δ _latfor horizontal feathering, δ _dlonfor longitudinal differential feathering.

4. one according to claim 3 the flight control method of tilting rotor depopulated helicopter transition section, it is characterized in that: after described nacelle inclination angle is more than 45 °, longitudinal mode of tilting rotor depopulated helicopter transition section is total apart from δ _cwith elevating rudder δ _ejointly control, horizontal side direction is aileron δ _awith yaw rudder δ _rjointly control; Then, the Longitudinal Control Law form of three phases is

X ₀(k)＝[*u *w *q u w q θ H] ^T，R(k)＝[u w q] ^T

In formula, δ _efor elevating rudder drift angle;

The horizontal side direction control law form of rear three phases is

X ₀(k)＝[*β *p *r β p r φ] ^T，R(k)＝[β p r] ^T。

In formula, δ _afor aileron drift angle, δ _rfor rudder.

5. any one according to Claims 1-4 the flight control method of tilting rotor depopulated helicopter transition section, it is characterized in that: the method for designing of inertial delay formula desalination apparatus is specially:

If the output of former control law is U _at (), the output of reconfigurable control rule is U _b(t), the output of switching law is U (t); Suppose at t ₀moment carries out control laws transformation, to t ₁moment, through exponential damping effect, and complete control laws transformation.The specific algorithm of this type desalination apparatus is

Parameter a is the principal element of impact desalination performance, and its value is larger, and fade time is shorter, otherwise its value is less, and fade time is longer.