Modeling and analyzing method for shimmy flexible tilting rotorcraft ship surface resonance

Abstract

The invention provides a shimmy flexible tilt rotorcraft ship surface resonance modeling and analysis method, which comprises the following steps: solving a landing gear vertical load static balance equation according to a right side half-spread kinetic model to obtain a wheel vertical load; according to the data in the first test curve and the vertical load of the airplane wheel, the compression amount of the airplane wheel and the buffer is obtained through interpolation; according to the compression amount of the airplane wheel and the buffer and the data of the second test curve, carrying out interpolation to obtain rigidity and damping coefficients of the airplane wheel and the buffer; according to the right side half-unfolding kinetic model and the rigidity and damping coefficients of the aircraft wheels and the buffers, a full-unfolding kinetic model of the aircraft is established; performing balancing analysis on rotor control to obtain linear mass, rigidity and a damping matrix of the fully-expanded kinetic model; and carrying out eigenvalue analysis on the linearized mass, rigidity and damping matrix to obtain a stability analysis result of the fully-expanded kinetic model.

Description

Technical Field

[0001] The present application belongs to the field of dynamic analysis technology, and in particular relates to a deck resonance modeling and analysis method for a swing-flexible tilt-rotor aircraft. Background Art

[0002] Tilt-rotor aircraft are new aircraft that have the advantages of vertical take-off and landing of helicopters and fast cruising and long range of fixed-wing propeller aircraft. The rotor of this type of aircraft is connected to the tilt nacelle on the wing. The nacelle has a tilt mechanism inside, which can make the rotor and the nacelle synchronously switch from horizontal position to vertical position. When the nacelle is tilted in the horizontal position, the rotor is equivalent to a traditional fixed-wing propeller, so that the aircraft has the ability to fly forward at high speed; when the nacelle is in the vertical position, the aircraft can imitate the vertical take-off and landing and hover in the air like a horizontal twin-rotor helicopter.

[0003] Compared with traditional single-rotor helicopters with tail rotors, tilt-rotor aircraft have many advantages such as long range and fast cruising, but their special features such as wing / rotor coupling and tilt-rotor nacelle will cause more complex aeroelastic stability problems. Summary of the invention

[0004] Purpose of the invention: The present invention proposes a dynamic modeling and analysis method for the deck resonance of a swing-flexible tiltrotor aircraft, in order to avoid and eliminate the deck resonance of the tiltrotor aircraft.

[0005] The present application provides a method for modeling and analyzing deck resonance of a swing-flexible tiltrotor aircraft, the method comprising:

[0006] Step 1: Establish the right half-span dynamics model of the aircraft;

[0007] Step 2: Solve the static balance equation of the vertical load of the landing gear according to the right half-extension dynamics model to obtain the vertical load of the wheel;

[0008] Step 3, interpolating the wheel and buffer compression according to the data in the first test curve and the vertical load of the wheel;

[0009] Step 4, interpolating the stiffness and damping coefficient of the wheel and the buffer according to the compression amount of the wheel and the buffer and the data of the second test curve;

[0010] Step 5: Establish a full-stretch dynamic model of the aircraft according to the right half-stretch dynamic model and the stiffness and damping coefficient of the wheel and the buffer;

[0011] Step 6: Perform a trim analysis on the rotor control to obtain the linearized mass, stiffness and damping matrix of the full-span dynamic model;

[0012] Step 7: Perform eigenvalue analysis on the linearized mass, stiffness and damping matrix to obtain the stability analysis result of the full-scale dynamic model.

[0013] Preferably, the step 1 specifically includes:

[0014] According to the structural characteristics of the right wing / rotor of the tiltrotor aircraft, based on the right-hand rule, a series of coordinate systems are established to describe the spatial positions of the center of mass of the fuselage, the center of mass of the wing root section, the center of mass of the wing end section, the center of mass of the tilt nacelle, the center of the hub, the center of mass of the universal hinge, and the center of mass of any section of the rotor.

[0015] Based on the coordinate system and Hamilton's principle, a right-side semi-expansion dynamics model is established.

[0016] Preferably, the step 2 specifically includes:

[0017] According to the position coordinates of the wheels in the right half-expansion dynamics model and the motion equations of the ship's roll, pitch and heave motions, a static balance equation of the landing gear vertical load is established;

[0018] Solve the static balance equation of the landing gear vertical load and obtain the vertical load on the wheel.

[0019] Preferably, the abscissa of the first test curve is the vertical load on the wheel, and the ordinate is the compression of the wheel and the buffer.

[0020] Preferably, the abscissa of the second test curve is the compression of the wheel and the buffer, and the ordinate is the stiffness and damping coefficient of the wheel and the buffer.

[0021] Preferably, the step 5 specifically includes:

[0022] According to the right half-span dynamics model and the stiffness and damping coefficient of the wheel and the buffer, a full-span dynamics model of the aircraft is established by adopting symmetric transformation.

[0023] Preferably, the step 6 specifically includes:

[0024] Under given state conditions, the rotor aerodynamic load is balanced and calculated so that the rotor thrust coefficient reaches a given value, and the rolling moment and pitching moment coefficients are both zero, thereby determining the rotor input collective pitch and cyclic pitch;

[0025] The input total pitch and periodic variable pitch are introduced into the full-span dynamics model, and the full-span dynamics model is linearized to obtain the linearized mass, stiffness and damping matrix of the full-span dynamics model.

[0026] Preferably, the stability analysis results include:

[0027] When the real part of the eigenvalue is positive, the full-scale dynamic model is unstable;

[0028] When the real part of the eigenvalue is negative, the full-scale dynamic model is stable.

[0029] Beneficial technical effects of this application:

[0030] The modeling and analysis method of deck resonance of a shimmy-flexible tiltrotor aircraft provided in this application fully considers the influence of ship motion on asymmetric load and deformation of landing gear, the influence of special mechanisms of tiltrotor aircraft such as tilt mechanism and universal joint on fuselage dynamics model, and the influence of different working states on rotor control balance. Under different balance states, the coupling relationship of each degree of freedom in the aeroelastic analysis model of the tiltrotor aircraft can be used to perform aeroelastic stability analysis on a shimmy-flexible tiltrotor aircraft under different sea conditions and different lift unloading. BRIEF DESCRIPTION OF THE DRAWINGS

[0031] Figure 1 It is a schematic diagram of the coordinate system of a full-span tiltrotor aircraft;

[0032] Figure 2 Schematic diagram of vertical force on landing gear;

[0033] Figure 3 Control trim calculation flow chart;

[0034] Figure 4 Flowchart of tiltrotor deck resonance stability analysis. DETAILED DESCRIPTION

[0035] Complex sea conditions put forward strict requirements on pilot control and helicopter deck resonance stability design. Therefore, it is necessary to fully consider the influence of fuselage rigid body displacement, wing / rotor elastic deformation and ship motion on wheel and buffer elastic deformation, establish a full-span dynamic model, conduct in-depth research on the tilt-rotor deck resonance stability analysis method, and form a complete set of deck resonance stability analysis technology to provide design means to avoid and eliminate tilt-rotor deck resonance.

[0036] The invention provides a dynamic modeling and analysis method for deck resonance of a swing-flexible tiltrotor aircraft. The tiltrotor aircraft structure consists of a rotor, a tilt nacelle, a wing and a body-landing gear system.

[0037] The specific steps include:

[0038] Step 1: Establish the right half-span dynamics model of the aircraft;

[0039] Step 2: Solve the static balance equation of the vertical load of the landing gear according to the right half-extension dynamics model to obtain the vertical load of the wheel;

[0040] Step 3, interpolating the wheel and buffer compression according to the data in the first test curve and the vertical load of the wheel;

[0041] Step 4, interpolating the stiffness and damping coefficient of the wheel and the buffer according to the compression amount of the wheel and the buffer and the data of the second test curve;

[0042] Step 5: Establish a full-stretch dynamic model of the aircraft according to the right half-stretch dynamic model and the stiffness and damping coefficient of the wheel and the buffer;

[0043] Step 6: Perform a trim analysis on the rotor control to obtain the linearized mass, stiffness and damping matrix of the full-span dynamic model;

[0044] Step 7: Perform eigenvalue analysis on the linearized mass, stiffness and damping matrix to obtain the stability analysis result of the full-scale dynamic model.

[0045] It should be noted that the full-span dynamics model fully considers the impact of ship motion on the asymmetric load of the landing gear and the left-right symmetric structural characteristics of the tilt-rotor fuselage.

[0046] When a tilt-rotor aircraft takes off and lands on a ship's deck, the ship is affected by the sea conditions and produces up and down heave, roll, and pitch motions. At this time, the ship will produce a vertical force on the landing gear, and the point of action is the landing point of the landing gear wheels on the ship's deck.

[0047] It is assumed that the inertial load on the landing point generated by the ship's rocking motion in a short period of time is a constant load, and the angle between the ship's deck plane and the sea level remains constant. Based on the above assumptions, the static balance equation of the landing gear vertical load is established, and the vertical load of the wheel is obtained by solving the static balance equation of the landing gear vertical load. Subsequently, based on the first test data, the compression of the wheel and the buffer is interpolated; based on the second test data, the compression of the wheel and the buffer, the stiffness and damping coefficient of the landing gear wheel and the buffer are interpolated.

[0048] In the process of modeling the deck resonance of the tiltrotor aircraft, the right half-span dynamic model of the aircraft is first established based on the right-handed coordinate system, and then the right half-span dynamic model of the aircraft is converted into the left half-span dynamic model of the aircraft using the left-handed coordinate system considering the influence of the left-right symmetrical distribution of the fuselage, and finally the two half-span dynamic models are merged into the full-span dynamic model of the aircraft. This modeling process will effectively improve the efficiency of the tiltrotor aircraft deck resonance modeling.

[0049] See also Figure 1-Figure 4 In other embodiments of the present application, the specific implementation steps of the present invention are as follows:

[0050] Step 1: Based on the structural characteristics of the right wing / rotor of the tiltrotor aircraft, a series of coordinate systems are established based on the right-hand rule to describe the spatial positions of the center of mass of the fuselage, the center of mass of the wing root section, the center of mass of the wing end section, the center of mass of the tiltrotor nacelle, the center of the hub, the center of mass of the universal hinge, and the center of mass of any section of the rotor. The coordinate system diagram is shown in Figure 1 .

[0051] Step 2: In the coordinate system, the kinetic energy and potential equations of the right wing / rotor of the tiltrotor aircraft, the potential energy equation of the tiltrotor nacelle, and the potential energy equation of the tiltrotor aircraft fuselage-landing gear are obtained, and then a right half-span dynamics model is established based on the Hamilton principle and the kinetic energy and potential energy equations.

[0052] Step 3: Consider the ship motion as simple harmonic motion, the ship pitches b jx 、Roll b jy and heave motionb jz The equation of motion.

[0053] b jx =b jxmax cos(ω 1 t)

[0054] b jy =b jymax cos(ω 2 t+Δ 1 )

[0055] b jz =b jzmax cos(ω 3 t+Δ 2 )

[0056] Among them, b jx,max 、b jy,max and b jz,max are the maximum roll, pitch angle and heave displacement respectively; ω 1 ,ω 2 and ω 3 are the angular frequencies of roll, pitch and heave respectively; Δ 1 and Δ 2 They are the phase difference between roll and pitch and the phase difference between roll and heave motion respectively.

[0057] Step 3: Perform a force analysis of the vertical load on the landing gear. The vertical force of the landing gear is as follows: Figure 2 As shown in the figure, the landing gear is subjected to the force P exerted on it by the ship surface N , a x 、a y and a z are the linear acceleration of the ship, ε x and ε yare the angular acceleration of the ship's motion, G is the body's gravity, and T is the body's lift. According to the position coordinates of the wheels in the right half-expansion dynamics model and the motion equations of the ship's roll, pitch, and heave motion, the static balance equation of the landing gear vertical load is established.

[0058]

[0059] P ML +P MR =Gcos(b jx )cos(b jy )-T+Ma z -P N

[0060]

[0061] Among them, (X N ,0,Z N )、(X m ,-Y m ,Z m ) and (X m ,Y m ,Z m ) are the coordinates of the front start wheel, main start left wheel and right wheel respectively.

[0062] Step 4: Solve the static balance equation of the vertical load of the landing gear to obtain the vertical load of the wheel.

[0063] Step 5: Based on the first test data and the polynomial interpolation method, an interpolation calculation is performed with the vertical load of the wheel as the horizontal coordinate, and the vertical coordinate obtained by interpolation is the compression amount of the wheel and the buffer.

[0064] Step 6: Based on the second test data and the polynomial interpolation method, interpolation calculation is performed with the compression of the wheel and the buffer as the horizontal axis, and the vertical axis obtained by interpolation is the stiffness and damping coefficient of the wheel and the buffer.

[0065] Step 7: Substitute the stiffness and damping coefficient of the wheel and the buffer into the right half-stretch dynamics model to obtain an updated right half-stretch dynamics model.

[0066] Step 8: Use the left-handed coordinate system to perform a symmetric transformation on the updated right half-expansion dynamics model to obtain the left half-expansion dynamics model.

[0067] Step 9: Merge the updated right half-stretch dynamics model and the left half-stretch dynamics model to obtain the full-stretch dynamics model of the aircraft. The full-stretch dynamics model has the following degrees of freedom.

[0068]

[0069] Among them, Xf , Y f and Z f Represents the three translational degrees of freedom of the body, θ fx ,θ fy and θ fz are the three rotational degrees of freedom of the body, v R 、w R and θ R They represent the linear displacement and axial torsion of the wing of the updated right half-span dynamic model, and is the flapping angle of the universal joint of the updated right half-span dynamic model, and They represent the linear displacements in three directions and the axial torsion of blade i in the updated right half-span dynamic model, v L 、w L and θ L They represent the linear displacement and axial torsion of the wing of the left half-span dynamic model respectively. and is the flapping angle of the universal joint of the half-span dynamic model on the left, and They respectively represent the linear displacement in three directions and the axial torsion of blade i in the left half-span dynamic model.

[0070] Step 10: Under given conditions, perform control and trim calculations on the rotor to make the rotor thrust T reach a given value and the rolling moment M x and pitch moment M y are all zero, thus determining the rotor input collective pitch θ 75 and periodic variable pitch θ 1c and θ 1s . See the control trim calculation flow chart for Figure 3 In the figure, Δθ represents the input total distance difference and periodic variable distance difference calculated in two iterative steps.

[0071] Step 11: Substitute the input total pitch and periodic variable pitch into the full-span dynamic model, and linearize the full-span dynamic model to obtain the linearized mass, stiffness and damping matrix of the full-span dynamic model (there are too many expressions and they are not listed separately).

[0072] Step 12: Perform eigenvalue analysis on the linearized mass, stiffness and damping matrix to obtain the stability analysis result of the full-scale dynamic model.

[0073] When the real part of the eigenvalue is positive, the full-scale dynamic model is unstable;

[0074] When the real part of the eigenvalue is negative, the full-scale dynamic model is stable.

[0075] The calculation and analysis process is as follows Figure 4 shown.

[0076] The present application provides a modeling and analysis method for deck resonance of a shimmy-flexible tilt-rotor aircraft. The aircraft model includes four parts: fuselage-landing gear system, wing, rotor and tilt nacelle. Considering the influence of the rigid body displacement of the fuselage, the elastic deformation of the wing / rotor and the ship motion on the elastic deformation of the landing gear, a full-span dynamic model of the aircraft is established. The aeroelastic stability analysis of the shimmy-flexible tilt-rotor aircraft under different sea conditions and different lift unloading can be performed under different trim states.

Claims

1. A method for modeling and analyzing deck resonance of a swing-flexible tiltrotor aircraft, characterized in that: The method comprises: Step 1: Establish the right half-span dynamics model of the aircraft; Step 2: Solve the static balance equation of the vertical load of the landing gear according to the right half-extension dynamics model to obtain the vertical load of the wheel; Step 3, interpolating the wheel and buffer compression according to the data in the first test curve and the vertical load of the wheel; Step 4, interpolating the stiffness and damping coefficient of the wheel and the buffer according to the compression amount of the wheel and the buffer and the data of the second test curve; Step 5: Establish a full-stretch dynamic model of the aircraft according to the right half-stretch dynamic model and the stiffness and damping coefficient of the wheel and the buffer; Step 6: Perform a trim analysis on the rotor control to obtain the linearized mass, stiffness and damping matrix of the full-span dynamic model; Step 7: Perform eigenvalue analysis on the linearized mass, stiffness and damping matrix to obtain the stability analysis result of the full-scale dynamic model.

2. The method according to claim 1, characterized in that The step 1 specifically includes: According to the structural characteristics of the right wing / rotor of the tiltrotor aircraft, based on the right-hand rule, a series of coordinate systems are established to describe the spatial positions of the center of mass of the fuselage, the center of mass of the wing root section, the center of mass of the wing end section, the center of mass of the tilt nacelle, the center of the hub, the center of mass of the universal hinge, and the center of mass of any section of the rotor. Based on the coordinate system and Hamilton's principle, a right-side semi-expansion dynamics model is established.

3. The method according to claim 2, characterized in that The step 2 specifically includes: According to the position coordinates of the wheels in the right half-expansion dynamics model and the motion equations of the ship's roll, pitch and heave motions, a static balance equation of the landing gear vertical load is established; Solve the static balance equation of the landing gear vertical load and obtain the vertical load on the wheel.

4. The method according to claim 3, characterized in that The abscissa of the first test curve is the vertical load on the wheel, and the ordinate is the compression of the wheel and the buffer.

5. The method according to claim 4, characterized in that The abscissa of the second test curve is the compression of the wheel and the buffer, and the ordinate is the stiffness and damping coefficient of the wheel and the buffer.

6. The method according to claim 5, characterized in that The step 5 specifically includes: According to the right half-span dynamics model and the stiffness and damping coefficient of the wheel and the buffer, a full-span dynamics model of the aircraft is established by adopting symmetric transformation.

7. The method according to claim 6, characterized in that The step 6 specifically includes: Under given state conditions, the rotor aerodynamic load is balanced and calculated so that the rotor thrust coefficient reaches a given value, and the rolling moment and pitching moment coefficients are both zero, thereby determining the rotor input collective pitch and cyclic pitch; The input total pitch and periodic variable pitch are introduced into the full-span dynamics model, and the full-span dynamics model is linearized to obtain the linearized mass, stiffness and damping matrix of the full-span dynamics model.

8. The method according to claim 7, characterized in that The stability analysis results include: When the real part of the eigenvalue is positive, the full-scale dynamic model is unstable; When the real part of the eigenvalue is negative, the full-scale dynamic model is stable.

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