CN112182753A - Control decoupling design method for tilt rotor helicopter - Google Patents

Control decoupling design method for tilt rotor helicopter Download PDF

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CN112182753A
CN112182753A CN202011020383.0A CN202011020383A CN112182753A CN 112182753 A CN112182753 A CN 112182753A CN 202011020383 A CN202011020383 A CN 202011020383A CN 112182753 A CN112182753 A CN 112182753A
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CN112182753B (en
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刘毅
苏小恒
刘宝方
孙强
马成江
万海明
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China Helicopter Research and Development Institute
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Abstract

The invention discloses a control decoupling design method of a tilt rotor helicopter, which comprises the following steps: establishing an initial longitudinal control strategy of the tilt rotor helicopter, establishing a pneumatic force and moment model and a gravity model of a component of the tilt rotor helicopter, and obtaining a nonlinear differential equation set based on a six-degree-of-freedom motion equation; according to the nonlinear differential equation set, selecting rotor deflection angles under interval values, solving the flight trim control quantity of the maximum speed of the tilt rotor helicopter and the required engine power, establishing a maximum trim speed trim result table under different rotor deflection angles, and determining a new longitudinal control strategy; determining a new flat flight speed trim result table based on the new longitudinal control strategy; selecting a rotor total pitch corresponding to the same steering rod longitudinal manipulation quantity under different rotor deflection angles from the table, and determining a decoupling coefficient between the rotor deflection angle and the rotor total pitch based on the rotor total pitch; and calculating the change of the collective pitch along with the inclination angle of the rotor according to the decoupling coefficient, thereby solving the total pitch of the rotor.

Description

Control decoupling design method for tilt rotor helicopter
Technical Field
The technical field of helicopters is related to, and specifically relates to a control decoupling design method for a tilt rotor helicopter.
Background
Tilt rotor helicopters have the handling characteristics of helicopters and fixed wing aircraft. When in helicopter mode, the rotor system generates lift force and pull force, and the longitudinal, transverse and course control is realized through the cyclic pitch change of the main rotor. The helicopter with the tilt rotor wing has the advantages that the lift force is generated by the wings in a fixed wing mode, the pulling force is generated by the rotor wing, the longitudinal, transverse and course control is realized through the elevator, the aileron and the rudder, and the degree of freedom of the variation of the pulling force azimuth angle of the rotor wing is increased relative to the helicopter by the helicopter with the tilt rotor wing. The tilt rotor helicopter is in a transition process from a helicopter mode to a fixed wing mode, a pilot mainly controls the tilt rotor to realize transition flight through a steering column longitudinally, a total pitch and a rotor deflection angle, but the control surface control under the variable pitch control and the fixed wing mode under the helicopter mode can be added into the control of the tilt rotor helicopter at the same time, so that the control surface redundancy phenomenon occurs. In addition, in the process of switching between the helicopter mode and the fixed wing mode, the control of the rotor deflection angle is strongly coupled with the control of the collective pitch and the longitudinal control, and the coupling causes that a pilot needs to carry out a large amount of control corrections with different amplitudes on each channel during transitional flight, increases the workload of the pilot and seriously influences the flight quality of the transitional flight.
Disclosure of Invention
The invention provides a longitudinal control strategy and a decoupling design method of a tilt rotor helicopter in the process of switching a helicopter mode and a fixed wing mode, and aims to solve the problems of redundancy of a control surface of the tilt rotor helicopter and coupling between control of a rotor deflection angle, total distance control and longitudinal control.
In order to realize the task, the invention adopts the following technical scheme:
a control decoupling design method of a tilt rotor helicopter comprises the following steps:
an initial longitudinal control strategy of the tilt rotor aircraft is formulated, and the initial longitudinal control strategy comprises the change of a longitudinal control rod, the tilt angle of a rotor wing and the change of the total pitch of the rotor wing along with the tilt angle of the rotor wing;
according to the established initial longitudinal control strategy, establishing a pneumatic force and moment model and a gravity model of the components of the tilt rotor helicopter; substituting the result obtained by calculation of the pneumatic force and moment model and the gravity model into a six-degree-of-freedom motion equation to obtain a nonlinear differential equation set;
according to the nonlinear differential equation set, selecting a rotor wing deflection angle under an interval value, and solving the flight trim manipulated variable of the maximum speed of the tilt rotor helicopter and the required engine power; the flight trim control quantity comprises a rotor total pitch, a steering column longitudinal control, a longitudinal periodic variable pitch and a total power, a maximum flat flight speed trim result table under different rotor deflection angles is established, and a new longitudinal control strategy is determined;
establishing a pneumatic force and moment model, a gravity model and a six-degree-of-freedom motion equation of a helicopter component based on a new longitudinal control strategy, determining a new nonlinear differential equation set from the pneumatic force and moment model and the gravity model, and then determining a new flat flight speed trim result table; selecting a rotor total pitch corresponding to the same steering rod longitudinal manipulation quantity under different rotor deflection angles from the table, and determining a decoupling coefficient between the rotor deflection angle and the rotor total pitch based on the rotor total pitch; and calculating the change of the collective pitch along with the inclination angle of the rotor according to the decoupling coefficient, so as to solve the total pitch of the rotor and finish the decoupling process.
Further, the method for determining the change of the longitudinal joystick comprises the following steps:
acquiring an initial operating range of the rotorcraft, wherein the initial operating range comprises an initial operating range [ a1, b1] of longitudinal cyclic variation and an initial operating range [ a2, b2] of the elevator, so that the transfer coefficient of the longitudinal cyclic variation is C1 (a1-b1)/N, and the transfer coefficient of the elevator is C2 (a2-b 2)/N; the variation of the longitudinal stick of the rotorcraft is set to (C1+ C2) × log.
Further, the determining a new longitudinal maneuver strategy includes:
according to the maximum flat flying speed trim result table, selecting a rotor deflection angle range of which the maximum speed is limited by longitudinal manipulation amount from the table, selecting a minimum rotor deflection angle from the table, and when the minimum rotor deflection angle is linearly reduced by straight 0 degrees, linearly reducing the transfer coefficient C1 of the longitudinal cyclic variable pitch to 0 along with the minimum rotor deflection angle, and generating a corresponding new longitudinal manipulation strategy once every change.
Further, the determining a decoupling factor between rotor deflection angle and rotor collective pitch includes:
and respectively carrying out linear fitting on different rotor wing deflection angles and different rotor wing total pitches, screening out all linear changing line segments from a fitting line of the rotor wing total pitches, and taking the slope of the line segment which is closest to the fitting line of the rotor wing deflection angles and has the minimum slope as the decoupling coefficient C3.
Further, the rotor collective pitchcolThe collective control amount + the collective pitch variation Δ col with the rotor pitch angle.
Further, the value of N is the total stroke of the joystick change.
A terminal device comprising a processor, a memory, and a computer program stored in the memory, the computer program, when executed by the processor, implementing the tilt-rotor helicopter operational decoupling design method steps.
A computer-readable storage medium having stored thereon a computer program which, when executed by a processor, performs the steps of the tilt-rotor helicopter maneuvering decoupling design method.
The invention has the following technical characteristics:
the invention solves the problem of redundant longitudinal control surfaces in the transition process of the tilt rotor helicopter, can solve the coupling phenomenon between the tilt angle and the total distance of the rotor in the tilt process of the rotor, reduces the operation load of a pilot and realizes the smooth transition of the tilt rotor helicopter.
Drawings
FIG. 1 is a schematic flow diagram of the process of the present invention;
FIG. 2 is a graph of C1 versus β;
FIG. 3 is a graph showing the variation of rotor collective pitch with rotor yaw angle β;
FIG. 4 shows the rotor collective pitch col for maximum speed at different rotor deflection angles β
FIG. 5 is a graph of collective control as a function of rotor deflection angle β during a transition from helicopter mode to fixed wing mode;
FIG. 6 is a graph of collective maneuvers as a function of speed during a transition from helicopter mode to fixed wing mode.
Detailed Description
The invention adopts a technical scheme that a design method of a longitudinal control strategy of a tilt rotor helicopter is shown in the attached drawing 1, and comprises the following steps:
step one, an initial longitudinal operation strategy of the tilt rotor aircraft is formulated, and the initial longitudinal operation strategy comprises the change of a longitudinal operation rod, the tilt angle of a rotor wing and the change of the total pitch of the rotor wing along with the tilt angle of the rotor wing.
Acquiring an initial operating range of the rotorcraft, wherein the initial operating range comprises an initial operating range [ a1, b1] of longitudinal cyclic variation and an initial operating range [ a2, b2] of the elevator, so that the transfer coefficient of the longitudinal cyclic variation is C1 (a1-b1)/N, and the transfer coefficient of the elevator is C2 (a2-b 2)/N; the value of N may be a total stroke of the joystick, which may be 100, for example.
The variation of the longitudinal stick of the rotorcraft is set to (C1+ C2) — log, where log represents the longitudinal stick amount; the rotor tilt angle is obtained from the overall design parameters of the rotorcraft and is a default configuration; the collective pitch is initially set to 0 as a function of rotor pitch.
In the transition mode, the helicopter and fixed wing modes of operation coexist, and as the nacelle pitch angle changes from the helicopter mode to the fixed wing mode, the longitudinal stick to longitudinal cyclic pitch transfer coefficient C1 decreases until it reaches 0. In order to obtain the change rule of the C1 along with the rotor wing deflection angle beta, the initial value of C1 is set to be 0.2 according to the range of longitudinal cyclic displacement xb of a certain tilt rotor helicopter, wherein the range is between-10 degrees and 10 degrees, and the transmission coefficient C2 from a longitudinal joystick to an elevator is a fixed value and does not change along with the change of the angle beta. The invention mainly explains a longitudinal control strategy and a decoupling design method in the transition process, so that a horizontal direction control strategy adopts default configuration.
xb＝0.2*log
TABLE 1 manipulation of channel assignments
Channel Range of Operating member
Longitudinal steering delta log -50％-50％ Longitudinal periodic variable pitch elevator
Total distance delta col 0°-36° Total pitch of rotor wing
Rotor deflection angle beta 0°-90° Rotor deflection angle
Step two, establishing a pneumatic force and moment model f of the tilt rotor helicopter component according to the established initial longitudinal control strategy1And a gravity model f2As shown in formula 1 and formula 2; substituting the result obtained by calculation of the pneumatic force and moment model and the gravity model into a six-degree-of-freedom motion equation
Figure BDA0002700423860000041
Obtaining a nonlinear differential equation set f3As shown in formula 3:
Figure BDA0002700423860000042
Figure BDA0002700423860000043
Figure BDA0002700423860000044
wherein u, v and w are components of a velocity vector on three coordinate axes of a body axis system, p, q and r are rotation angular velocities on the three coordinate axes of the body axis system, theta, phi and sigma are three attitude angles, and Ix、Iy、Iz、Ixy、Ixz、IyzIs moment of inertia, wherein, IxIs moment of inertia on the X axis, IxyThe moment of inertia in the XY direction is represented,logfor longitudinal manipulation amount,latIs a transverse operation quantity,pedIs the pedal operation quantity,colFor rotor gross pitch, β is rotor deflection angle, and a and B are coefficient matrices derived from submodels of the individual components.
Selecting rotor deflection angles beta and taking M as intervals according to the nonlinear differential equation set, and solving the flight trim manipulated variable of the maximum speed of the tilt rotor helicopter and the required engine power by using a Newton iteration method; the flight trim control quantity comprises a rotor total pitch, a steering rod longitudinal control, a longitudinal periodic variable pitch and a total power, and a maximum flat flight speed trim result table under different rotor deflection angles is established.
According to a maximum flat flight speed trim result table, selecting a rotor deflection angle range of which the maximum speed is limited by longitudinal control amount from the table, selecting a minimum rotor deflection angle from the table, and when the minimum rotor deflection angle is linearly reduced by straight 0 degrees, linearly reducing the transfer coefficient C1 of longitudinal cyclic pitch change to 0 along with the minimum rotor deflection angle, generating a corresponding new longitudinal control strategy every time the minimum rotor deflection angle is changed, wherein each longitudinal control strategy comprises the change of a longitudinal control rod, the change of a total pitch along with a rotor inclination angle and a rotor tilt angle; wherein the rotor gross pitch still adopts the default configuration of step 1 along with rotor inclination and rotor tilt angle, and the rotor gross pitch is set to be 0 along with the change of rotor inclination.
And (4) a maximum flat flying speed balancing result table under different rotor wing deflection angles. As shown in table 2; wherein M is 2 ° to 10 °, for example 5 °.
TABLE 2 maximum flat flight velocity trim results for different rotor deflection angles
Figure BDA0002700423860000051
Figure BDA0002700423860000061
The maximum flying speed of the tilt rotor helicopter with the rotor deflection angle between 90 and 35 degrees is mainly limited by the longitudinal manipulation amount through the comparison of the maximum flying speed balancing results of different rotor deflection angles in the table 2. When the rotor deflection angle is between 35 degrees and 0 degrees, the maximum flight speed of the tilt rotor helicopter is mainly limited by the power of an engine, and the longitudinal cyclic pitch change is linearly changed along with the rotor deflection angle. According to the principle of maximizing the flight performance of the tilt rotor helicopter, the transfer coefficient C1 from the longitudinal joystick to the longitudinal cyclic pitch change cycle is linearly reduced from beta equal to 35 degrees until C1 equal to 0 degrees when beta equal to 0 degrees, as shown in FIG. 2.
Step four, replacing the initial longitudinal control strategy in the step two with the longitudinal control strategy determined in the step three, establishing a pneumatic force and moment model, a gravity model and a six-degree-of-freedom motion equation of the helicopter component according to the same method in the step two, determining a new nonlinear differential equation set according to the pneumatic force and moment model and the gravity model, and determining a new flat flight speed balancing result table through the step three; selecting a rotor total pitch corresponding to the same steering column longitudinal manipulation quantity log under different rotor deflection angles beta in the table, determining a decoupling coefficient C3 between the rotor deflection angle and the rotor total pitch based on the rotor total pitch, and calculating the change delta col of the total pitch along with the rotor inclination angle according to the decoupling coefficient:
Δcol＝C3*Δβ
where Δ β represents the amount of change in rotor pitch.
The specific determination method comprises the following steps:
and respectively carrying out linear fitting on different rotor wing deflection angles and different rotor wing total pitches, screening out all linear changing line segments from a fitting line of the rotor wing total pitches, and taking the slope of the line segment which is closest to the fitting line of the rotor wing deflection angles and has the minimum slope as the decoupling coefficient C3.
As shown in table 3, three different longitudinal maneuvers were chosen for comparison, since the longitudinal maneuver was smaller at smaller rotor yaw angles. As can be seen from table 3 and fig. 3, when the rotor deflection angle β is greater than 35 ° and smaller than 35 °, the rotor collective pitch and the rotor deflection angle β respectively change approximately linearly with two different slopes, and in order to obtain a better operation efficiency, a slope with a smaller slope is selected as a decoupling coefficient, and finally, a decoupling coefficient C3 between the deflection angle and the collective pitch can be obtained. The red frame in fig. 4 is the relationship between the total pitch manipulation range and the rotor deflection angle after considering the linkage relationship between the total pitch and the rotor deflection angle, and the black point in the frame is the total pitch trim amount of different speed points under different rotor deflection angles, and it can be seen from the figure that all the corresponding relationships between the rotor deflection angle and the total pitch manipulation fall within the manipulation limit range between the rotor deflection angle and the total pitch manipulation; the determined C3 in this example is 0.21.
Δcol＝0.21*Δβ
TABLE 3 trim results for the same longitudinal manipulated variable for different rotor yaw angles
Figure BDA0002700423860000071
Figure BDA0002700423860000081
And (3) verification process:
solving the total pitch of the rotor wing by utilizing the total pitch obtained by calculation in the step four and the change delta col of the rotor wing inclination anglecolThe calculation formula is as follows:
rotor wing total oarDistance betweencolTotal pitch manipulated variable + change in total pitch with rotor angle Δ col formula 4
To obtain the total pitch of the rotorcolThen the total pitch of the rotorcolThe rotor wing deflection angle is taken into a nonlinear differential equation set, the rotor wing deflection angles are changed at intervals of 2-3 degrees, the longitudinal manipulated variable, the transverse manipulated variable, the pedal manipulated variable, the total pitch and the speed of each rotor wing deflection angle under rated power are solved, and then the total pitch manipulated variable is obtained through calculation of a formula 4; based on the total distance manipulated variable, a comparison curve of the total distance manipulated variable of the driver when the decoupling retroversion rotary-wing helicopter and the decoupling retroversion rotary-wing helicopter are in equal power transition is solved, and as can be seen from the graph in fig. 5 and 6, the relationship of the manipulation linkage is formed after the decoupling, and the total distance manipulation load is obviously reduced in the transitional flight process. And after a flight control system of the tilt rotor helicopter is updated, the actual flight of a driver is carried out, and the coupling relation between the rotor deflection angle and the total distance operation and between the rotor deflection angle and the longitudinal operation is verified.
The above embodiments are only used for illustrating the technical solutions of the present application, and not for limiting the same; although the present application has been described in detail with reference to the foregoing embodiments, it should be understood by those of ordinary skill in the art that: the technical solutions described in the foregoing embodiments may still be modified, or some technical features may be equally replaced; such modifications and substitutions do not substantially depart from the spirit and scope of the embodiments of the present application, and are intended to be included within the scope of the present application.

Claims (8)

1. A control decoupling design method of a tilt rotor helicopter comprises the following steps:
an initial longitudinal control strategy of the tilt rotor aircraft is formulated, and the initial longitudinal control strategy comprises the change of a longitudinal control rod, the tilt angle of a rotor wing and the change of the total pitch of the rotor wing along with the tilt angle of the rotor wing;
according to the established initial longitudinal control strategy, establishing a pneumatic force and moment model and a gravity model of the components of the tilt rotor helicopter; substituting the result obtained by calculation of the pneumatic force and moment model and the gravity model into a six-degree-of-freedom motion equation to obtain a nonlinear differential equation set;
according to the nonlinear differential equation set, selecting a rotor wing deflection angle under an interval value, and solving the flight trim manipulated variable of the maximum speed of the tilt rotor helicopter and the required engine power; the flight trim control quantity comprises a rotor total pitch, a steering column longitudinal control, a longitudinal periodic variable pitch and a total power, a maximum flat flight speed trim result table under different rotor deflection angles is established, and a new longitudinal control strategy is determined;
establishing a pneumatic force and moment model, a gravity model and a six-degree-of-freedom motion equation of a helicopter component based on a new longitudinal control strategy, determining a new nonlinear differential equation set from the pneumatic force and moment model and the gravity model, and then determining a new flat flight speed trim result table; selecting a rotor total pitch corresponding to the same steering rod longitudinal manipulation quantity under different rotor deflection angles from the table, and determining a decoupling coefficient between the rotor deflection angle and the rotor total pitch based on the rotor total pitch; and calculating the change of the collective pitch along with the inclination angle of the rotor according to the decoupling coefficient, so as to solve the total pitch of the rotor and finish the decoupling process.
2. The tilt rotor helicopter maneuvering decoupling design method of claim 1, wherein the change in the longitudinal stick is determined by:
acquiring an initial operating range of the rotorcraft, wherein the initial operating range comprises an initial operating range [ a1, b1] of longitudinal cyclic variation and an initial operating range [ a2, b2] of the elevator, so that the transfer coefficient of the longitudinal cyclic variation is C1 (a1-b1)/N, and the transfer coefficient of the elevator is C2 (a2-b 2)/N; the variation of the longitudinal stick of the rotorcraft is set to (C1+ C2) × log.
3. The tilt-rotor helicopter maneuvering decoupling design method of claim 1, wherein the determining a new longitudinal maneuvering strategy comprises:
according to the maximum flat flying speed trim result table, selecting a rotor deflection angle range of which the maximum speed is limited by longitudinal manipulation amount from the table, selecting a minimum rotor deflection angle from the table, and when the minimum rotor deflection angle is linearly reduced by straight 0 degrees, linearly reducing the transfer coefficient C1 of the longitudinal cyclic variable pitch to 0 along with the minimum rotor deflection angle, and generating a corresponding new longitudinal manipulation strategy once every change.
4. The tilt-rotor helicopter maneuvering decoupling design method of claim 1, wherein said determining a decoupling factor between rotor deflection angle and rotor collective pitch comprises:
and respectively carrying out linear fitting on different rotor wing deflection angles and different rotor wing total pitches, screening out all linear changing line segments from a fitting line of the rotor wing total pitches, and taking the slope of the line segment which is closest to the fitting line of the rotor wing deflection angles and has the minimum slope as the decoupling coefficient C3.
5. The tilt-rotor helicopter maneuvering decoupling design method of claim 1, wherein the collective pitch of the rotors iscolThe collective control amount + the collective pitch variation Δ col with the rotor pitch angle.
6. The tilt rotor helicopter maneuvering decoupling design method of claim 1, wherein the value of N is the total travel of the change in the maneuvering lever.
7. A terminal device comprising a processor, a memory and a computer program stored in the memory, characterized in that the computer program, when executed by the processor, carries out the steps of the method according to any one of claims 1-6.
8. A computer-readable storage medium, in which a computer program is stored which, when being executed by a processor, carries out the steps of the method according to any one of claims 1 to 6.
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CN113076601A (en) * 2021-04-20 2021-07-06 中国直升机设计研究所 Helicopter slope take-off and landing calculation and test flight method
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CN114323551A (en) * 2022-03-15 2022-04-12 中国空气动力研究与发展中心低速空气动力研究所 Tilting transition corridor wind tunnel experiment balancing method and system for tilting rotorcraft

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