Distributed multi-propeller tilt-rotor aircraft and transient transition aeroelastic analysis method

Abstract

The invention provides a distributed multi-propeller tilt rotor aircraft and a transient transition aeroelastic analysis method. The method comprises the following steps: step 1, establishing a coordinate system of a distributed multi-propeller / tilt wing configuration coupling system; 2, establishing a coupling system structure model according to the coordinate system; 3, establishing a coupled system aerodynamic model according to the coordinate system; 4, wake bending correction is carried out on the aerodynamic model of the coupling system; 5, according to the structure model and the aerodynamic model of the distributed multi-propeller / tilt wing configuration coupling system, forming a dynamic overall array of the coupling system; 6, the tilt transition corridor parameters are substituted into the dynamic overall array, and a transient dynamic equation of the coupling system is obtained; and 7, carrying out numerical solution on the transient kinetic equation by adopting a Newmark time finite element method, and solving the aeroelastic response of the transient transition process of the distributed multi-paddle / tilt wing configuration coupling system under the condition of gradually updating the tilt transition parameters and modal parameters.

Description

Technical Field

[0001] The present application relates to the field of helicopter rotor dynamics, and in particular to a distributed multi-blade tilt-rotor aircraft and a transient transition aeroelastic analysis method. Background Art

[0002] Distributed multi-blade tilt-rotor aircraft combines the vertical take-off and landing capabilities of helicopters with the high-speed cruising capabilities of fixed-wing aircraft. It is one of the important configurations for the electrification, intelligence, high-speed and long-range development of future rotorcraft. Distributed multi-rotors and large aspect ratio tilt-rotor wings are its main lift systems. During transition flight, the rotor aerodynamic force interacts with the elastic deformation of the large aspect ratio wing. The dynamic wake interaction of the rotor / wing, the change of the rotor / wing coupling dynamic characteristics over time, and the mutual excitation of the rotor / wing aerodynamic and inertial loads through the structure produce strong self-excited oscillations, increase the transient response, and may induce structural damage, resulting in complex rotor / wing coupling aeroelastic dynamics problems. At present, the rotor / wing aeroelastic coupling mechanism under transient transition conditions is not fully understood, and there is still a lack of effective analysis methods for the unsteady and nonlinear aeroelastic response characteristics caused by rotor pull reversal, wing stall at large angles of attack, and flight attitude and control transients. Summary of the invention

[0003] The present application provides a distributed multi-blade tilt-rotor aircraft and a transient transition aeroelastic analysis method, which can better predict the unsteady aeroelastic behavior characteristics of the distributed multi-blade tilt-rotor aircraft during the transient transition process.

[0004] In a first aspect, the present application provides a method for transient transition aeroelastic analysis of a distributed multi-propeller tilt-rotor aircraft, the method comprising:

[0005] Step 1: Establish a coordinate system for the distributed multi-propeller / tilt-wing configuration coupling system;

[0006] Step 2: According to the coordinate system, a distributed multi-propeller / tilt-wing configuration coupling system structure model is established;

[0007] Step 3: Establishing an aerodynamic model of a distributed multi-propeller / tilt-wing configuration coupling system according to the coordinate system;

[0008] Step 4: Perform wake bending correction on the aerodynamic model of the distributed multi-propeller / tilt-wing configuration coupling system;

[0009] Step 5: Based on the structural model and aerodynamic model of the distributed multi-rotor / tilt-wing configuration coupling system, a dynamic overall array of the distributed multi-rotor / tilt-wing configuration coupling system is formed;

[0010] Step 6: Substitute the tilt-rotation transition corridor parameters into the overall dynamics matrix to obtain the transient dynamic equation of the distributed multi-rotor / tilt-rotation wing configuration coupling system;

[0011] Step 7: Use the Newmark time finite element method to numerically solve the transient dynamic equations, and obtain the aeroelastic response of the system's transient transition process by gradually updating the tilting transition parameters and modal parameters.

[0012] Specifically, step 1 includes:

[0013] Step 11: Establish a ground fixed inertial coordinate system with the connection point between the wing root and the fuselage as the origin;

[0014] Step 12: According to the ground fixed inertial coordinate system, a coordinate system of the distributed multi-propeller / tilt-wing configuration coupling system is established, wherein the coordinate system includes the wing undeformed coordinate system X I Y I Z I , wing deformation coordinate system X W Y W Z W , rotor nacelle coordinate system X P Y P Z P , the hub does not rotate the coordinate system X H Y H Z H , hub rotation coordinate system X R Y R Z R , hub plane coordinate system X U Y U Z U , the blade undeformed coordinate system X B Y B Z B and blade deformation coordinate system

[0015] Specifically, the transformation matrices satisfy:

[0016] T UI =T UR T RH T HP T PW T WI ;

[0017] T PI =T PW T WI ;

[0018] T UD =T UR T RB T BD .

[0019] Step 13: Calculate the transformation matrix T of each coordinate system in the coordinate systemWI , T PW , T HP , T RH , T UR , T RB , T BD .

[0020] Specifically, step 2 includes:

[0021] Step 21: Construct the system strain energy δU of the blade b :

[0022]

[0023] Step 22: Construct the system strain energy δU of the wing W :

[0024]

[0025] Step 23: Construct the system kinetic energy δT of the blade b :

[0026]

[0027] Step 24: Construct the System Kinetic Energy δT of the Wing w :

[0028]

[0029] Step 25: Construct the system kinetic energy δT of the nacelle p :

[0030]

[0031] Specifically, step 3 includes:

[0032] Step 31: Construct the aerodynamic virtual work δW of the blade b :

[0033] Among them, L u is the spanwise aerodynamic force of the blade, L v is the blade chord-wise aerodynamic force, L w is the vertical aerodynamic force of the blade, is the aerodynamic moment at the aerodynamic center of the blade;

[0034] Step 32: Construct the aerodynamic virtual work δW of the wing W :

[0035]

[0036] Specifically, step 4 includes:

[0037] Using the modified formula to calculate the aerodynamic virtual work δW W To make corrections:

[0038]

[0039] Where: α is the wake inclination angle, K R is the wake bending correction coefficient, take K R =1;κ s is the wake curvature in the rotor rolling direction, κ c is the curvature of the wake in the rotor pitch direction.

[0040] Specifically, the overall dynamics matrix of the distributed multi-rotor / tilt-wing configuration coupling system is:

[0041]

[0042] in: ξ b They are respectively the acceleration degree of freedom, velocity degree of freedom and displacement degree of freedom of the blade in the non-rotating coordinate system; x h They are the acceleration degree of freedom, velocity degree of freedom and displacement degree of freedom of the connection point between the wing and the nacelle respectively; x ww are the acceleration degree of freedom, velocity degree of freedom and displacement degree of freedom on the wing except the connection point between the wing and the nacelle; M bb is the blade mass matrix, C bb is the blade damping matrix, K bb is the blade stiffness matrix, F b is the blade load vector; M bxh is the blade / wing coupling mass matrix, C bxh is the blade / wing coupling damping matrix, K bxh is the blade / wing coupling stiffness matrix; M xhb is the wing / blade coupling mass matrix, C xhb is the wing / blade coupling damping matrix, K xhb is the wing / blade coupling stiffness matrix; M xhxh is the mass matrix of the wing / nacelle connection point, C xhxh is the wing / nacelle connection point damping matrix, K xhxh is the wing / nacelle connection point stiffness matrix, F xh is the load vector at the wing / nacelle connection point; M wxh 、M xhw 、M ww is the mass matrix of the wing except the connection point between the wing and the nacelle, C wxh , C xhw , C wwis the damping matrix of the wing points except the connection point between the wing and the nacelle, K wxh , K xhw , K ww is the stiffness matrix of the wing except the connection point between the wing and the nacelle, F w is the load vector at the wing / nacelle connection point.

[0043] Specifically, the tilt transition corridor parameters include the forward flight speed vf of the aircraft, the wing tilt angle wp, the tilt angular velocity dwp, the tilt angular acceleration ddwp, and the rotor total pitch and cyclic pitch parameters during the tilt transition process.

[0044] In a second aspect, the present application provides a distributed multi-blade tilt-rotor aircraft, which is implemented using the distributed multi-blade tilt-rotor aircraft transient transition aeroelastic analysis method as described above.

[0045] In summary, this patent provides a method for transient transition aeroelastic analysis of a distributed multi-blade tilt-rotor aircraft. Aiming at the transient transition aeroelastic behavior characteristics of a multi-blade tilt-rotor aircraft, this patent develops a transient transition aeroelastic analysis method for a rotor / wing flexible multi-body coupling system, reveals the unsteady aeroelastic behavior characteristics of the tilt-rotor transition state, and provides a theoretical basis and technical support for the development of this type of aircraft model. BRIEF DESCRIPTION OF THE DRAWINGS

[0046] Figure 1 A schematic diagram of the vertical bending response of the wing tip provided for this application;

[0047] Figure 2 Schematic diagram of wing lift coefficient provided for this application;

[0048] Figure 3 Schematic diagram of the chord-wise bending response of the wing tip provided for this application

[0049] Figure 4 Schematic diagram of wing drag coefficient provided for this application;

[0050] Figure 5 A schematic diagram of the rotor flapping taper angle provided for this application;

[0051] Figure 6 Schematic diagram of the longitudinal swing angle provided for this application;

[0052] Figure 7 Schematic diagram of chord-wise flapping angle provided for this application. DETAILED DESCRIPTION

[0053] The purpose of the present invention is to provide a distributed rotor / tilt-wing aeroelastic coupling analysis method under transient transition mode, to reveal the formation mechanism of aeroelastic instability, to grasp the influence law of parameters, and to assist in the technical development and model development of distributed multi-propeller tilt-rotor aircraft.

[0054] Embodiment 1

[0055] Distributed multi-propeller / tilt-wing coupled dynamic equations, based on the Lagrangian coordinate system, this paper describes the configuration of any point on the coupled system in the inertial coordinate system relative to the inertial coordinate system, and uses the Hamilton principle to derive the multi-propeller / tilt-wing coupled dynamic equations, the expressions of which are as follows.

[0056]

[0057] In the above formula: δU, system strain energy; δT, system kinetic energy; δW, system external force virtual work, each part is composed as follows:

[0058]

[0059] In the above formula: subscripts W, P, and b represent wing, nacelle, and blade respectively, N b 、N r are the number of blades and rotors respectively. The nacelle is assumed to be a rigid body, and the influence of its support stiffness connected with the wing and rotor is not considered. Its aerodynamic force is ignored and only its kinetic energy is considered.

[0060] The present application provides a method for transient transition aeroelastic analysis of a distributed multi-propeller tilt-rotor aircraft, the method comprising:

[0061] Step 1: Establish a coordinate system for the distributed multi-propeller / tilt-wing configuration coupling system;

[0062] Specifically, step 1 includes:

[0063] Step 11: Establish a ground fixed inertial coordinate system with the connection point between the wing root and the fuselage as the origin;

[0064] Step 12: According to the ground fixed inertial coordinate system, a coordinate system of the distributed multi-propeller / tilt-wing configuration coupling system is established, wherein the coordinate system includes the wing undeformed coordinate system X I Y I Z I , wing deformation coordinate system X W Y W Z W , rotor nacelle coordinate system X P Y P Z P , the hub does not rotate the coordinate system X H Y H Z H, hub rotation coordinate system X R Y R Z R , hub plane coordinate system X U Y U Z U , the blade undeformed coordinate system X B Y B Z B and blade deformation coordinate system

[0065] Step 13: Calculate the transformation matrix T of each coordinate system in the coordinate system WI , T PW , T HP , T RH , T UR , T RB , T BD ;

[0066] Among them, the transformation matrix T WI is the transformation matrix from the wing deformation coordinate system to the wing undeformed coordinate system, the transformation matrix T PW is the transformation matrix from the rotor nacelle coordinate system to the wing deformation coordinate system, the transformation matrix T HP is the transformation matrix from the hub non-rotating coordinate system to the rotor nacelle coordinate system, the transformation matrix T RH The transformation matrix T is the transformation matrix from the hub rotating coordinate system to the hub non-rotating coordinate system. UR is the transformation matrix from the hub plane coordinate system to the hub rotation coordinate system, the transformation matrix T RB is the transformation matrix from the blade undeformed coordinate system to the hub plane coordinate system, the transformation matrix T BD is the transformation matrix for transforming the blade deformation coordinate system into the blade undeformed coordinate system;

[0067] Specifically, the transformation matrices satisfy:

[0068] T UI =T UR T RH T HP T PW T WI ;

[0069] T PI =T PW T WI ;

[0070] T UD =T UR T RB T BD .

[0071] Step 2: According to the coordinate system, a distributed multi-propeller / tilt-wing configuration coupling system structure model is established;

[0072] Specifically, step 2 includes:

[0073] Step 21: Construct the system strain energy δU of the blade b :

[0074]

[0075] Step 22: Construct the system strain energy δU of the wing W :

[0076]

[0077] Step 23: Construct the system kinetic energy δT of the blade b :

[0078]

[0079] Step 24: Construct the System Kinetic Energy δT of the Wing w :

[0080]

[0081] Step 25: Construct the system kinetic energy δT of the nacelle p :

[0082]

[0083] It should be noted that in the structural modeling of the distributed multi-propeller / tilt-wing configuration coupling system, both the blades and the wings can be analyzed using the medium deformation beam model, and the nacelle can be regarded as a rigid body.

[0084] In the above formula: σ, ε are the stress and strain of the blade or wing section respectively; η, ζ are the coordinates of any point on the section in the section coordinate system respectively, L represents the rotor radius or the half span of the wing, strain σ, ε are calculated by Hooke's law and deformation coordination relationship respectively; u, v, w, φ are the degrees of freedom of the blade, x is h ,y h 、z h ,φ h , α h , h is the degree of freedom at the connection point P; u w 、v w 、w w ,φ w is the wing degree of freedom;

[0085] Step 3: Establishing an aerodynamic model of a distributed multi-propeller / tilt-wing configuration coupling system according to the coordinate system;

[0086] Specifically, step 3 includes:

[0087] Step 31: Construct the aerodynamic virtual work δW of the blade b :

[0088] Among them, L u is the spanwise aerodynamic force of the blade, L v is the blade chord-wise aerodynamic force, L w is the vertical aerodynamic force of the blade, is the aerodynamic moment of the aerodynamic center of the blade.

[0089] Step 32: Construct the aerodynamic virtual work δW of the wing W :

[0090]

[0091] It should be noted that the expression of the aerodynamic virtual work of the wing is the same as that of the blade, but due to the unique configuration of the distributed multi-propeller / tilt-wing aircraft, part of the wing is always in the rotor wake. This section of the wing is called the slipstream area, so that the chord-wise velocity of the wing in the slipstream area has an increment brought by the rotor wake, and the shrinkage formula of the rotor wake radius in the slipstream area is:

[0092]

[0093] In the above formula: R, C T are the rotor radius and the thrust coefficient respectively, h is the nacelle length, R h is the radius of the slipstream zone.

[0094] The chordwise velocity increment v1 of the wing in the slipstream region satisfies the gas continuity equation, and the following relationship exists:

[0095]

[0096] In the above formula: v i is the mean value of rotor induced inflow, U Tw is the velocity component of the incoming flow velocity in the chordwise direction of the wing.

[0097] Step 4: Perform wake bending correction on the aerodynamic model of the distributed multi-propeller / tilt-wing configuration coupling system;

[0098] Specifically, step 4 includes: using the modified formula to calculate the aerodynamic virtual work δW W To make corrections:

[0099]

[0100] Where: α is the wake inclination angle, K R is the wake bending correction coefficient, take K R =1;κs is the wake curvature in the rotor rolling direction, κ c is the curvature of the wake in the rotor pitch direction.

[0101] Specifically, using the formula Calculate the curvature of the wake in the direction of the rotor roll; use the formula Calculate the curvature of the rotor's wake in pitch direction.

[0102] In the above formula: p h ,q h are the tilt angular rates of the rotor axis in the roll and pitch directions, respectively, where the rotor axis pitch angular rate is the wing tilt angular rate, It is the horizontal and vertical flapping angular velocity of the rotor blade tip trajectory plane.

[0103] It should be noted that by modifying K R The values ​​may include or ignore the effects of rotor wake bending.

[0104] It should be noted that when calculating the dynamic inflow with wake bending correction taken into account, during the tilt transition flight of the distributed multi-rotor / tilt-wing configuration, the wing drives the rotor to tilt at a large angle, and the upper rotor is similar to the maneuvering flight state of a helicopter. The rotor wake geometry will produce a wake distortion effect. The dynamic inflow model needs to consider the impact of the rotor wake bending on the inflow and add a wake bending correction factor to the inflow "gain" matrix.

[0105] Step 5: Based on the structural model and aerodynamic model of the distributed multi-rotor / tilt-wing configuration coupling system, a dynamic overall array of the distributed multi-rotor / tilt-wing configuration coupling system is formed;

[0106] Specifically, the overall dynamics matrix of the distributed multi-rotor / tilt-wing configuration coupling system is:

[0107]

[0108] in: ξ b They are respectively the acceleration degree of freedom, velocity degree of freedom and displacement degree of freedom of the blade in the non-rotating coordinate system;

[0109] x h They are the acceleration degree of freedom, velocity degree of freedom and displacement degree of freedom of the connection point between the wing and the nacelle respectively;

[0110] x ww They are the acceleration degree of freedom, velocity degree of freedom and displacement degree of freedom on the wing except the connection point between the wing and the nacelle;

[0111] M bbis the blade mass matrix, C bb is the blade damping matrix, K bb is the blade stiffness matrix, F b is the blade load vector;

[0112] M bxh is the blade / wing coupling mass matrix, C bxh is the blade / wing coupling damping matrix, K bxh is the blade / wing coupling stiffness matrix;

[0113] M xhb is the wing / blade coupling mass matrix, C xhb is the wing / blade coupling damping matrix, K xhb is the wing / blade coupling stiffness matrix;

[0114] M xhxh is the mass matrix of the wing / nacelle connection point, C xhxh is the wing / nacelle connection point damping matrix, K xhxh is the wing / nacelle connection point stiffness matrix, F xh is the wing / nacelle connection point load vector;

[0115] M wxh 、M xhw 、M ww is the mass matrix of the wing except the connection point between the wing and the nacelle, C wxh , C xhw , C ww is the damping matrix of the wing points except the connection point between the wing and the nacelle, K wxh , K xhw , K ww is the stiffness matrix of the wing except the connection point between the wing and the nacelle, F w is the load vector at the wing / nacelle connection point.

[0116] It should be noted that, according to Hamilton's principle, the present application groups the blades into rotor equations through multi-blade coordinate transformation, and then couples them with the nacelle and the wing to form the overall dynamics matrix of the distributed multi-rotor / tilt-wing configuration coupling system.

[0117] The overall system matrix can be written as

[0118]

[0119] In the above formula: M S (t), C S (t), K S (t) is the system structural modal mass, damping, and stiffness matrix;

[0120] q(t), is the system generalized coordinate vector; F(t) is the force that is independent of the response, is the force related to the response (including structural nonlinear terms and aerodynamic terms).

[0121] Step 6: Substitute the tilt-rotation transition corridor parameters into the overall dynamics matrix to obtain the transient dynamic equations of the distributed multi-rotor / tilt-wing configuration coupling system.

[0122] Among them, the parameters of the tilt transition corridor include the forward flight speed vf of the aircraft, the wing tilt angle wp, the tilt angular velocity dwp, the tilt angular acceleration ddwp, and the rotor total pitch and cyclic pitch parameters during the tilt transition process.

[0123] Step 7: Use the Newmark time finite element method to numerically solve the transient dynamic equations, and obtain the aeroelastic response of the system's transient transition process by gradually updating the tilting transition parameters and modal parameters.

[0124] It should be noted that by analyzing the calculation results of the aeroelastic response, the unsteady aeroelastic behavior characteristics of the tilt-rotor transition state can be revealed, the influence laws of parameters can be mastered, and the technical development and model development of distributed multi-propeller tilt-rotor aircraft can be assisted.

[0125] In addition, the present application provides a distributed multi-blade tilt-rotor aircraft, and the distributed multi-blade tilt-rotor aircraft is implemented using the distributed multi-blade tilt-rotor aircraft transient transition aeroelastic analysis method as described above.

[0126] In summary, the aeroelastic analysis method for transient transition of distributed multi-rotor aircraft developed in this patent takes into account the aeroelastic coupling between the distributed multi-rotor, nacelle and large aspect ratio flexible wing, and accounts for the influence of the rotor wake on the wing and the influence of the rotor wake bending on the rotor inflow during the tilt transition process. It better predicts the unsteady aeroelastic behavior characteristics of the transient transition process of the distributed multi-rotor aircraft, and masters the aeroelastic coupling mechanism of the distributed multi-rotor / tilt-wing under transient transition conditions, providing an efficient and accurate analysis method for the dynamic design of this type of aircraft.

[0127] Embodiment 2

[0128] like Figure 1-7 As shown, the embodiment of the present application takes the transient tilt transition process of a distributed multi-propeller tilt-rotor aircraft as an example to describe in detail a method for transient transition aeroelastic analysis of a distributed multi-propeller tilt-rotor aircraft provided by the present application:

[0129] The tilt transition process of the tilt-wing is from helicopter mode to fixed-wing mode. The wing starts to tilt from t=1s and ends at t=5.11s. The wing tilts linearly, and the tilt angle rate is a constant: 18° / s. The wing tilt angle in helicopter mode (the angle between the wing chord direction and the vertical direction) is 10°, and the wing rigid attack angle in fixed-wing mode is 6° (wing tilt angle: 84°).

[0130] like Figure 1 , Figure 2 and Figure 3 , the lift coefficient of the wing during the tilt transition process can approximately reflect the trend of the vertical force change of the wing. By checking whether the vertical flapping trend of the wing is consistent with the lift coefficient change during the tilt transition process, it can be judged that the self-aerodynamic lift is dominant in the vertical load. The chord-wise response analysis is similar to the vertical one. The wing drag coefficient decreases during the tilt process, and the chord-wise (backward) deformation of the wing decreases. In the fixed-wing mode, the wing aerodynamic drag is small and is pulled forward by the rotor. The chord-wise deformation tends to 0 and reaches a steady-state value. The wing tilt angle before the tilt transition is 10°. At this time, the rotor is in the forward flight state of the helicopter. The asymmetric inflow of the rotor disc leads to the flapping modal response of the rotor βc and βs. Since the wing tip (rotor base) is in periodic motion, the rotor modal response is a periodic oscillation. In the transition flight mode, the axial flow of the rotor increases and the flapping taper angle decreases. The aerodynamic environment of the rotor disc plane becomes more complex, the motion coupling between the rotor and the wing intensifies, and the periodic flapping modal oscillation becomes stronger. At the end of the tilting, the wing tilt angle is 84°, the rotor is in axial flow flight, the asymmetric phenomenon of the rotor disc inflow basically disappears, the rotor βc and βs modal responses eventually approach 0, and the blade stabilizes at the taper flapping angle. At this time, the rotor thrust also tends to be constant (axial flow) and no longer changes periodically, the load transmitted to the wing end by the rotor tends to be constant, and the wing tip response also reaches a steady-state value.

Claims

1. A transient transition aeroelastic analysis method for a distributed multi-blade tilt-rotor aircraft, characterized in that: The method comprises: Step 1: Establish a coordinate system for the distributed multi-propeller / tilt-wing configuration coupling system; Step 2: According to the coordinate system, a distributed multi-propeller / tilt-wing configuration coupling system structure model is established; Step 3: Establishing an aerodynamic model of a distributed multi-propeller / tilt-wing configuration coupling system according to the coordinate system; Step 4: Perform wake bending correction on the aerodynamic model of the distributed multi-propeller / tilt-wing configuration coupling system; Step 5: Based on the structural model and aerodynamic model of the distributed multi-rotor / tilt-wing configuration coupling system, a dynamic overall array of the distributed multi-rotor / tilt-wing configuration coupling system is formed; Step 6: Substitute the tilt-rotation transition corridor parameters into the overall dynamics matrix to obtain the transient dynamic equation of the distributed multi-rotor / tilt-rotation wing configuration coupling system; Step 7: Use the Newmark time finite element method to numerically solve the transient dynamic equations, and obtain the aeroelastic response of the system's transient transition process by gradually updating the tilting transition parameters and modal parameters.

2. The method according to claim 1, characterized in that Step 1 includes: Step 11: Establish a ground fixed inertial coordinate system with the connection point between the wing root and the fuselage as the origin; Step 12: According to the ground fixed inertial coordinate system, a coordinate system of the distributed multi-propeller / tilt-wing configuration coupling system is established, wherein the coordinate system includes the wing undeformed coordinate system X I Y I Z I , wing deformation coordinate system X W Y W Z W , rotor nacelle coordinate system X P Y P Z P , the hub does not rotate the coordinate system X H Y H Z H , hub rotation coordinate system X R Y R Z R , hub plane coordinate system X U Y U Z U , the blade undeformed coordinate system X B Y B Z B and blade deformation coordinate system Step 13: Calculate the transformation matrix T of each coordinate system in the coordinate system WI 、T PW 、T HP 、T RH 、T UR 、T RB 、T BD . The transformation matrices satisfy: T UI =T UR T RH T HP T PW T WI ; T PI =T PW T WI ; T UD =T UR T RB T BD 。 3. The method according to claim 1, characterized in that Step 2 includes: Step 21: Construct the system strain energy δU of the blade b : Step 22: Construct the system strain energy δU of the wing W : Step 23: Construct the System Kinetic Energy δT of the Blade b : Step 24: Construct the System Kinetic Energy δT of the Wing w : Step 25: Construct the system kinetic energy δT of the nacelle p :

4. The method according to claim 1, characterized in that: Step 3 includes: Step 31: Construct the aerodynamic virtual work δW of the blade b : Among them, L u is the spanwise aerodynamic force of the blade, L v is the blade chord-wise aerodynamic force, L w is the vertical aerodynamic force of the blade, is the aerodynamic moment of the blade aerodynamic center; Step 32: Construct the aerodynamic virtual work δW of the wing W :

5. The method according to claim 1, characterized in that Step 4 includes: Using the modified formula to calculate the aerodynamic virtual work δW b To make corrections: Where: α is the wake inclination angle, K R is the wake bending correction coefficient, take K R =1;κ s is the wake curvature in the rotor rolling direction, κ c is the curvature of the wake in the rotor pitch direction.

6. The method according to claim 1, characterized in that The overall dynamics matrix of the distributed multi-rotor / tilt-wing configuration coupling system is: in: ξ b They are respectively the acceleration degree of freedom, velocity degree of freedom and displacement degree of freedom of the blade in the non-rotating coordinate system; x h They are the acceleration degree of freedom, velocity degree of freedom and displacement degree of freedom of the connection point between the wing and the nacelle respectively; x ww are the acceleration degree of freedom, velocity degree of freedom and displacement degree of freedom on the wing except the connection point between the wing and the nacelle; M bb is the blade mass matrix, C bb is the blade damping matrix, K bb is the blade stiffness matrix, F b is the blade load vector; M bxh is the blade / wing coupling mass matrix, C bxh is the blade / wing coupling damping matrix, K bxh is the blade / wing coupling stiffness matrix; M xhb is the wing / blade coupling mass matrix, C xhb is the wing / blade coupling damping matrix, K xhb is the wing / blade coupling stiffness matrix; M xhxh is the mass matrix of the wing / nacelle connection point, C xhxh is the wing / nacelle connection point damping matrix, K xhxh is the wing / nacelle connection point stiffness matrix, F xh is the load vector at the wing / nacelle connection point; M wxh 、M xhw 、M ww is the mass matrix of the wing except the connection point between the wing and the nacelle, C wxh , C xhw , C ww is the damping matrix of the wing points except the connection point between the wing and the nacelle, K wxh , K xhw , K ww is the stiffness matrix of the wing except the connection point between the wing and the nacelle, F w is the load vector at the wing / nacelle connection point.

7. The method according to claim 1, characterized in that The parameters of the tilt transition corridor include the forward flight speed vf of the aircraft, the wing tilt angle wp, the tilt angular velocity dwp, the tilt angular acceleration ddwp, and the rotor total pitch and cyclic pitch parameters during the tilt transition process.

8. According to the method of claim 1, the transient dynamic equation is numerically solved by using the Newmark time finite element method, and the aeroelastic response of the system transient transition process is obtained by gradually updating the tilting transition parameters and modal parameters.

9. A distributed multi-propeller tilt-rotor aircraft, characterized in that: The distributed multi-blade tilt-rotor aircraft is realized by using the distributed multi-blade tilt-rotor aircraft transient transition aeroelastic analysis method as described in any one of claims 1 to 8.

Images