CN104460681B - The flight control method of tilting rotor depopulated helicopter changeover portion - Google Patents

Abstract

A kind of flight control method of tilting rotor depopulated helicopter changeover portion, it is characterized in that, this method is using Optimum-preview control by the constraint of the emerging system equation of motion, foreseeable perfect condition and optimal performance index constraint information, and the flight to tilting rotor depopulated helicopter changeover portion is controlled；Specific method is：Nonlinear model is set up to tilting rotor depopulated helicopter changeover portion first, then trim linearisation obtains the linear equation of state space description under different nacelle inclination angles；Secondly, based on linear model, flying speed, height and the attitude angle in Optimum-preview control device, control transient process are designed；Finally formed for tilting rotor depopulated helicopter between different nacelle inclination angles and seamlessly transit Inertia of design delayed type desalination apparatus.It is an object of the invention to provide a kind of flight control method of tilting rotor depopulated helicopter changeover portion, the information such as foreseeable desired speed, height and attitude angle are merged, the handoff-security between tilting rotor depopulated helicopter mode is completed.

Description

The flight control method of tilting rotor depopulated helicopter changeover portion

Technical field

The present invention relates to the flight control method of tilting rotor depopulated helicopter changeover portion, more particularly to a kind of optimal prediction The flight control method of control, belongs to technical field of flight control.

Background technology

Tiltrotor is a kind of mixing aircraft, preceding as fixed wing aircraft can both be flown, while can also be as straight The machine of liter realizes VTOL and hovering like that.The property of tiltrotor, extends it in dual-use middle use model Enclose." Eagle Eye " are a tilting rotor depopulated helicopters currently developed in the U.S..

The problem of tiltrotor brings many new and technological difficulties：On the one hand, aircraft control mode and offline mode More complicated, the model for adding changeover portion describes difficulty, and existing airvane control has pulling force vector controlled again, do not gone straight up to only , there is the handoff procedure between helicopter mode and fixed-wing pattern again in machine pattern, fixed-wing pattern.On the other hand, in pattern Aerodynamic Characteristics are extremely complex in handoff procedure, and engine nacelle has stronger pitching couple when verting and bowed / plunging motion coupling is faced upward, Aerodynamic intenference is serious, and aircraft stability is poor.Therefore, aircraft how is made to stably accomplish tilting rotor Switching between machine pattern is even more important for safe flight.And requirement phase of the tilting rotor depopulated helicopter to flight control system It is more harsher than someone tiltrotor.

For the research of the research of tiltrotor changeover portion, particularly changeover portion control program, domestic and foreign scholars are proposed A variety of control method.In addition to most classical control methods, Dr.Anthony J.Calise and Rolf T.Rysdyk Adaptive nonlinear control method is proposed, this method uses network and models coupling.However, these control methods are often only sharp With the past and present information of system.But for aircraft control, Future Trajectory information is to realize planning and design.In advance See that control not only considers the current information of system, current control program is determined always according to the Future Information of system, therefore can To improve the control performance of system.

At present, it is therefore foreseen that control has been successfully applied in some fields., can be using prediction control during such as mechanical device active damping Control action is just added before vibration is produced or before interference signal appearance, control energy is reduced a lot；And for example in lathe, machine In the path following control of people etc., because path is mostly known in advance, track following error can be greatly reduced using prediction control. And it is seldom in the application of flying vehicles control technical field.

The content of the invention

Technical problem to be solved：

It is an object of the invention to provide a kind of flight control method of tilting rotor depopulated helicopter changeover portion, fusion can be pre- The information such as the desired trajectory and desired speed seen, complete the handoff-security between tilting rotor depopulated helicopter mode.

Technical scheme：

In order to realize function above, the invention provides a kind of flight controlling party of tilting rotor depopulated helicopter changeover portion Method, it is characterised in that this method is constrained using Optimum-preview control by the emerging system equation of motion, foreseeable perfect condition And optimal performance index constraint information, the flight to tilting rotor depopulated helicopter changeover portion is controlled；Specific method is：

Nonlinear model is set up to tilting rotor depopulated helicopter changeover portion first, then the trim under different nacelle inclination angles Linearisation obtains the linear equation of state space description；

Secondly, based on linear model, design flying speed in Optimum-preview control device, control transient process, height and Attitude angle；

Finally formed for tilting rotor depopulated helicopter between different nacelle inclination angles and seamlessly transit Inertia of design delayed type Desalination apparatus.

Concretely, the different nacelle angled states are β_M0 °:15 °, 15 °:30 °, 30 °:45 °, 45 °:60 °, 60°:75 °, 75 °:90°；In β_MFor 15 °, 30 °, 45 °, 60 °, 75 °, 90 °, the rotation of verting corresponding to above-mentioned six stages is set up The linear model of wing depopulated helicopter changeover portion, be expressed as { A1, B1, C1, D1 }, { A2, B2, C2, D2 } ..., A6, B6,C6,D6}}；Wherein, above-mentioned Ai, Bi, Ci, Di represent the sytem matrix of linear model respectively, each stage Ai, Bi, Ci, Di } include longitudinal model { Ai_lon,Bi_lon,Ci_lon,Di_lonAnd horizontal lateral model { Ai_lat,Bi_lat,Ci_lat,Di_lat}。

As a kind of preferred, when the nacelle inclination angle is no more than 45 °, the longitudinal direction of tilting rotor depopulated helicopter changeover portion Mode is always away from δ_cWith longitudinal feathering δ_lonJointly control, it is horizontal lateral for horizontal feathering δ_latBecome with the longitudinal differential cycle Away from δ_dlonJointly control；Then the Longitudinal Control Law form of preceding three phases is

X₀(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, q For rate of pitch, before variable plus * represents the difference of the variable measurements and ideal value；δ_cFor always away from δ_lonBecome for longitudinal cycle Away from θ is the angle of pitch；H is flying height；

Laterally control law form is the horizontal stroke of preceding three phases

X₀(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 before roll angle, variable plus * represents this The difference of variable measurements and ideal value；δ_latFor horizontal feathering, δ_dlonFor longitudinal differential feathering

As a kind of preferred, after the nacelle inclination angle is more than 45 °, the longitudinal direction side of tilting rotor depopulated helicopter changeover portion Formula is always away from δ_cWith elevator δ_eJointly control, horizontal is laterally aileron δ_aWith rudder δ_rJointly control；Three phases is vertical after then It is to control law form

X₀(k)=[^*u ^*w ^*q u w q θ H]^T, R (k)=[u w q]^T

In formula, δ_eFor lifting angle of rudder reflection；

The lateral control law form of the horizontal stroke of three phases is afterwards

X₀(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 former control law is output as U_A(t), reconfigurable control rule is output as U_B(t), switching law is output as U (t)；Assuming that in t₀Moment is controlled rule switching, to t₁At the moment, acted on by exponential damping,AndComplete control laws transformation.The specific algorithm of the type desalination apparatus is

Parameter a is the principal element of influence desalination performance, and its value is bigger, and fade time is shorter, conversely, its value is smaller, it is light The change time is longer.

Beneficial effect：

It is an object of the invention to provide a kind of flight control method of tilting rotor depopulated helicopter changeover portion, fusion can be pre- Desired speed, height and the attitude angle information seen, complete the handoff-security between tilting rotor depopulated helicopter mode.

A kind of flight control method of tilting rotor depopulated helicopter changeover portion of the present invention, is controlled using optimal prediction System passes through the letter such as emerging system equation of motion hard constraint, foreseeable perfect condition soft-constraint and optimal performance index soft-constraint Breath, the flight to tilting rotor depopulated helicopter changeover portion is controlled.

Brief description of the drawings

The present invention is further described with reference to the accompanying drawings and examples：

Fig. 1 is tilting rotor depopulated helicopter changeover portion speed safe range；

Fig. 2 is the longitudinal Optimum-preview control device of three phases before tilting rotor depopulated helicopter changeover portion of the present invention Structure；

Fig. 3 is control laws transformation desalination apparatus simulation curve figure of the present invention；

Fig. 4 is that the longitudinal velocity response at 15 degree of the nacelle inclination angle obtained using flight control method of the present invention is bent Line；

Fig. 5 is that the longitudinal velocity response at 45 degree of the nacelle inclination angle obtained using flight control method of the present invention is bent Line；

Fig. 6 is that the longitudinal velocity response at 60 degree of the nacelle inclination angle obtained using flight control method of the present invention is bent Line；

Fig. 7 is the one-dimensional track emulation curve obtained using flight control method of the present invention；

Fig. 8 is the two-dimentional track emulation curve obtained using flight control method of the present invention.

Embodiment

The present invention provides a kind of optimal prediction flight control method of tilting rotor depopulated helicopter changeover portion, to send out this Bright purpose, technical scheme and effect are clearer, clearly, and referring to the drawings and give an actual example to the present invention further specifically It is bright.It should be appreciated that specific implementation described herein is not intended to limit the present invention only to explain the present invention.

Using the present invention tilting rotor depopulated helicopter changeover portion optimal prediction flight control method, specifically according to Lower step.

Step 1, the nonlinear model for initially setting up tilting rotor depopulated helicopter changeover portion.Under body axis system, fly Row device dynamics and kinematics six degree of freedom equation group can use following form.

In computing power and torque, the effect of each section such as fuselage, rotor, wing, horizontal tail and vertical fin is considered.Such as machine The composition of the side force in body axle y-axis direction is as follows

Y=Y_fuse+Y_wing+Y_ROT+Y_HT+Y_VT

In formula, Y_fuseIt is the side force that fuselage is produced, Y_wingIt is the side force that wing is produced, Y_ROTIt is the side force that rotor is produced, Y_HT It is the side force that horizontal tail is produced, Y_VTIt is the side force that vertical fin is produced.

Fuselage, rotor, wing, horizontal tail and vertical fin are carried out after mathematical modeling respectively, built in MATLAB/Simulink The nonlinear mathematical model of tilting rotor depopulated helicopter.

Then the trim under different nacelle angles to system is realized using the trim functions in MATLAB/Simulink, then , will be non-near with flat spot using trim input quantity and trim condition amount as the parameter of linmod functions using linmod functions Linear model is linearized.Because the motion of tilting rotor depopulated helicopter changeover portion is with very strong time-varying characteristics, by changeover portion Six stages are divided, nacelle angle of inclination beta is corresponded to respectively_M0 °:15 °, 15 °:30 °, 30 °:45 °, 45 °:60 °, 60 °:75 °, 75 °: 90°.In β_MFor 15 °, 30 °, 45 °, 60 °, 75 °, 90 ° are set up the tilting rotor depopulated helicopter mistake for corresponding to above-mentioned six stages The linear model of section is crossed, is expressed as { { A1, B1, C1, D1 }, { A2, B2, C2, D2 } ..., { A6, B6, C6, D6 } }.Wherein { Ai, Bi, Ci, the Di } in each stage is included with x=[u w q θ H]^TFor state variable, with U=[δ_c δ_lon δ_e] it is control Longitudinal model { Ai of amount_lon,Bi_lon,Ci_lon,Di_lonAnd with x=[β p r φ]^TFor state variable, with U=[δ_lat δ_dlon δ_a δ_r] for the lateral model { Ai of horizontal stroke of controlled quentity controlled variable_lat,Bi_lat,Ci_lat,Di_lat}。

Understand technical scheme for the ease of the public, illustrated physical quantity of the present invention with table 1 below：

Table 1

Step 2, based on above-mentioned six linear models, design Optimum-preview control device.

Control program is determined first, it is as follows：

(1) when being no more than 45 ° at nacelle inclination angle, longitudinal mode of tilting rotor depopulated helicopter changeover portion is always away from δ_cWith Longitudinal feathering δ_lonJointly control, it is horizontal lateral for horizontal feathering δ_latWith longitudinal differential feathering δ_dlonJoint control System.

(2) after nacelle inclination angle is more than 45 °, longitudinal mode of tilting rotor depopulated helicopter changeover portion is always away from δ_cAnd lifting Rudder δ_eJointly control, horizontal is laterally aileron δ_aWith rudder δ_rJointly control.

According to above-mentioned control program, the state equation in six stages in step 1 is divided, preceding three phases are obtained Lengthwise movement state equation be

The horizontal lateral movement state equation of preceding three phases is

The lengthwise movement state equation of three phases is afterwards

The horizontal lateral movement state equation of three phases is afterwards

Above-mentioned linear model is carried out with certain sampling time by discretization.Obtain the linear discrete time that following formula is represented System.

X (k+1)=Ax (k)+BU (k)

Y (k)=Cx (k)

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

E (k)=R (k)-y (k)

Then following error signal system is exported

Or

X₀(k+1)=Φ X₀(k)+GΔU(k)+G_RΔR(k+1)

For error system, it is assumed that it is known its since present moment until following M_RLetter is evaluated in the desired value of step, definition Number is the following quadratic form comprising error term and input item：

Wherein：Q is positive semidefinite matrix；Q_e, H is positive definite matrix.Q and H are the weight matrix determined by designer.

There are several solutions for solving the optimal control problem of error system, conventional one kind is optimized according to partial differential Method, this method is the form for first assuming control input, is substituted into evaluation function, requires that this function takes minimum value according to partial differential, So as to obtain prediction feed-forward coefficients.

If the control input form of Optimum-preview control is

Here, F_R(j) it is coefficient undetermined.It is using above method solving result

F₀=-[H+G^TPG]^-1G^TPΦ

F_R(j)=- [H+G^TPG]^-1G^T(ξ^T)^j-1PG_R (j≥1)

(ξ=Φ+GF₀)

P=Q+ Φ^TPΦ-Φ^TPG[H+G^TPG]^-1G^TPΦ

If making F₀=[F_e,F_x], system initial value is zero, is obtained

It is the model for making aircraft forward flight speed within the stage reach safe transition in the purpose of each phase controller Enclose, i.e., in safe range as shown in Figure 1.It is with reference to the controller architecture of controlling plan design

The Longitudinal Control Law form of preceding three phases is

X₀(k)=[^*u ^*w ^*q u w q θ H]^T, R (k)=[u w q]^T

The longitudinal controller structure chart of preceding three phases is as shown in Figure 2.

Laterally control law form is the horizontal stroke of preceding three phases

X₀(k)=[^*β ^*p ^*r β p r φ]^T, R (k)=[β p r]^T

The Longitudinal Control Law form of three phases is afterwards

X₀(k)=[^*u ^*w ^*q u w q θ H]^T, R (k)=[u w q]^T

The lateral control law form of the horizontal stroke of three phases is afterwards

X₀(k)=[^*β ^*p ^*r β p r φ]^T, R (k)=[β p r]^T

Realize and seamlessly transit between step 3, six phase controllers, inertial delay formula desalination apparatus is devised for this, if former Control law is output as U_A(t), reconfigurable control rule is output as U_B(t), switching law is output as U (t).Assuming that in t₀When Rule switching is controlled quarter, to t₁At the moment, acted on by exponential damping,And Complete control laws transformation.The specific algorithm of the type desalination apparatus is

Parameter a is the principal element of influence desalination performance, and its value is bigger, and fade time is shorter, conversely, its value is smaller, it is light The change time is longer.

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

In order to verify validity of the present invention in the flight control of tilting rotor changeover portion, following emulation experiment is carried out.It is imitative True instrument uses MATLAB softwares, and object is analyzed using XV-15 tiltrotor aircrafts.Simulation parameter is set such as the institute of table 2 Show.

Table 2

With β in 0~5s_MTrim linear model at=15 ° is controlled, with β in 5~15s_MTrim at=30 ° Linear model is controlled, with β in 15~55s_MTrim linear model at=60 ° is controlled, forward flight speed u in table 2 Setting safety fall in the safe range shown in Fig. 1.The tracking mode of speed is set in preceding 15s, u is with acceleration a_u= 2m/s², w is with acceleration a_w=0.12m/s²Climb；The step signal that amplitude is 10 is inputted in 15s moment u, w inputs amplitude is 2.95 step signal.If elemental height is 100m, according to above-mentioned steps design control Optimum-preview control device, control effect is such as Shown in Fig. 4 to Fig. 8, wherein Fig. 7, all stage track shown in 8 calculate according to Fig. 4 to Fig. 6 rate curve and obtained.In Fig. 4 Ordinate to Fig. 6 represents that present speed subtracts the difference of initial velocity, in Fig. 4, although stable state mistake occurs in u control result Difference, but in the range of safe transition, do not interfere with the transition of tilting rotor.There is a decline when w control is initial in figure 6 Process, but be due to that w initial values are more than 0, so occurring that the height in the short time declines phenomenon, climb mode can be recovered afterwards, From the change of Fig. 7 altitude curve it is also seen that.The purpose controlled in this emulation essentially consists in speed control, be not track with Track, therefore track can have deviation with ideal trajectory.Above-mentioned simulation result fully shows, based on the tilting rotor for most having prediction to control Unmanned helicopter system has good response performance.

It is understood that for those of ordinary skills, can be with technique according to the invention scheme and its hair Bright design is subject to equivalent or change, and all these changes or replacement should all belong to the guarantor of appended claims of the invention Protect scope.

Claims (1)

1. a kind of flight control method of tilting rotor depopulated helicopter changeover portion, it is characterised in that this method is using optimal pre- Control is seen by the constraint of the emerging system equation of motion, foreseeable perfect condition and optimal performance index constraint information, to inclining The flight of switch rotor depopulated helicopter changeover portion is controlled；Specific method is：

Nonlinear model is set up to tilting rotor depopulated helicopter changeover portion first, then trim is linear under different nacelle inclination angles Change the linear equation for obtaining state space description；

Secondly, based on linear model, flying speed, height and the attitude in Optimum-preview control device, control transient process are designed Angle；Finally formed for tilting rotor depopulated helicopter between different nacelle inclination angles and seamlessly transit the desalination of Inertia of design delayed type Device；

The different nacelle angled states are β_M0 °:15 °, 15 °:30 °, 30 °:45 °, 45 °:60 °, 60 °:75 °, 75 °:90°； In β_MFor 15 °, 30 °, 45 °, 60 °, 75 °, 90 °, the tilting rotor depopulated helicopter transition corresponding to above-mentioned six stages is set up The linear model of section, is expressed as { { A1, B1, C1, D1 }, { A2, B2, C2, D2 } ..., { A6, B6, C6, D6 } }；Wherein, Ai, Bi, Ci, Di represent the sytem matrix of linear model respectively, and { Ai, Bi, Ci, the Di } in each stage includes longitudinal model {Ai_lon,Bi_lon,Ci_lon,Di_lonAnd horizontal lateral model { Ai_lat,Bi_lat,Ci_lat,Di_lat}；

When the nacelle inclination angle is no more than 45 °, longitudinal mode of tilting rotor depopulated helicopter changeover portion is always away from δ_cAnd longitudinal direction Feathering δ_lonJointly control, it is horizontal lateral for horizontal feathering δ_latWith longitudinal differential feathering δ_dlonJointly control；Then The Longitudinal Control Law form of preceding three phases is

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

X₀(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 to bow * is added to represent the difference of the variable measurements and ideal value before elevation angle speed, variable；δ_cFor always away from δ_lonFor longitudinal feathering, θ is The angle of pitch；H is flying height；

Laterally control law form is the horizontal stroke of preceding three phases

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

X₀(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 before roll angle, variable plus * represents the variable The difference of measured value and ideal value；δ_latFor horizontal feathering, δ_dlonFor longitudinal differential feathering；

After the nacelle inclination angle is more than 45 °, longitudinal mode of tilting rotor depopulated helicopter changeover portion is always away from δ_cAnd elevator δ_eJointly control, horizontal is laterally aileron δ_aWith rudder δ_rJointly control；The Longitudinal Control Law form of three phases is after then

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

X₀(k)=[* u * w * q u w q θ H]^T, R (k)=[u w q]^T

In formula, δ_eFor lifting angle of rudder reflection；

The lateral control law form of the horizontal stroke of three phases is afterwards

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

X₀(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；

The method for designing of inertial delay formula desalination apparatus is specially：

If former control law is output as U_A(t), reconfigurable control rule is output as U_B(t), switching law is output as U (t)；In t₀When Rule switching is controlled quarter, to t₁At the moment, acted on by exponential damping,And Control laws transformation is completed, the specific algorithm of the desalination apparatus is

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

Parameter a is the principal element of influence desalination performance, and its value is bigger, and fade time is shorter, conversely, its value is smaller, during desalination Between it is longer.