Transient unsteady rotor aerodynamic load calculation method and device

Abstract

The invention provides a transient unsteady rotor aerodynamic load calculation method and device. The method comprises the following steps: generating geometric characteristics of a blade grid of a rotor system, generating an initial wake of a rotor, and calculating an initial flow field of the rotor; updating a geometric position and a flight state parameter of a rotor blade surface grid under the current time step, and calculating a rotor wake structure and rotor flow field information of the current time step; calculating blade aerodynamic force, constructing a rotor kinetic equation to solve aeroelastic response, judging whether the aeroelastic response is converged or not, and updating a rotor wake structure and rotor flow field information; calculating the load of the current time step; and judging the size relationship between the load calculation time and the set calculation time, and outputting the rotor wake structure, the rotor flow field, the blade aerodynamic force, the aeroelastic response and the load information in the whole calculation process. By means of the method, time domain flow field information can be accurately constructed in the continuous change process of the flight state, manipulation input and the like of the helicopter, and the transient rotor aerodynamic load at any moment is obtained through solving.

Description

Technical Field

[0001] This invention belongs to the field of helicopter rotor aeroelastic dynamics, and particularly relates to a method and device for calculating transient unsteady rotor aerodynamic loads. Background Technology

[0002] During helicopter flight, rotor blades vibrate and deform under the influence of aerodynamic forces and centrifugal forces. This blade deformation also affects the calculation of aerodynamic loads on the blade profile, creating an aeroelastic coupling effect. Transient unsteady rotor aerodynamic load calculation methods can consider the transient aerodynamic changes of the blades under flapping, flaring, and pitch-changing motions. Furthermore, they can capture the transient aerodynamic characteristics of the rotor in different states such as hovering, forward flight, and side flight, making the simulation input more consistent with actual flight control and improving the reliability of the simulation results.

[0003] Currently, computational fluid dynamics (CFD) is the primary method for calculating transient rotor aerodynamic loads. This method discretizes the fluid control equations (such as the Navier-Stokes equations), introduces a time term to solve for unsteady flow fields, and thus obtains the transient aerodynamic loads on the upper and lower surfaces of the blade. It can account for the effects of complex geometry and multiphysics coupling, resulting in high computational accuracy. However, CFD methods require refined time steps for unsteady calculations, typically necessitating lengthy numerical iterations to meet convergence conditions. Furthermore, high-precision simulations result in an extremely large number of meshes, leading to an exponential increase in computational load and demanding hardware requirements. Moreover, CFD methods are highly dependent on the model; general turbulence models contain empirical parameter assumptions, resulting in insufficient accuracy for simulating complex unsteady flows. Experimental data is often required for correction during modeling and analysis. Summary of the Invention

[0004] To address the problems of existing CFD methods requiring refined time steps and lengthy numerical iterations to meet convergence conditions for unsteady calculations, resulting in an exponential increase in computational cost, and strong model dependence requiring experimental data for correction during modeling and analysis, this invention provides a method for calculating transient unsteady rotor aerodynamic loads. This method accurately constructs time-domain flow field information during continuously changing helicopter flight conditions and control inputs, and solves for transient rotor aerodynamic loads at any given time. The technical solution is as follows: Firstly, a method for calculating transient unsteady rotor aerodynamic loads is provided, the method comprising: Acquire rotor system parameters and initialize them. Generate rotor blade mesh geometry features. Based on the rotor blade surface mesh geometry position and flight state parameters at the initial time, generate the rotor initial wake and calculate the rotor initial flow field. Update the rotor blade surface mesh geometry position and flight state parameters at the current time step, and calculate the rotor wake structure and rotor flow field information at the current time step. Calculate the aerodynamic force of the blade and construct the rotor dynamics equation to solve the aeroelastic response. Determine whether the aeroelastic response has converged, and update the rotor wake structure and rotor flow field information until the blade aeroelastic response converges. Calculate the load at the current time step; determine the relationship between the load calculation time and the set calculation time, until the overall rotor wake calculation is completed, and output the rotor wake structure, rotor flow field, blade aerodynamic force, aeroelastic response and load information of the entire calculation process.

[0005] Optionally, the method specifically includes: Step 1: Obtain the rotor system parameters and complete the initialization of the rotor system parameters, including rotor configuration, geometric parameters, and dynamic parameters; Step 2: Based on the rotor configuration and geometric parameters, generate the rotor system blade mesh geometric features, and read the rotor blade surface mesh geometric position and flight state parameters at the initial moment; Step 3: Based on the initial geometric position of the rotor blade surface grid and flight state parameters, generate the initial rotor wake and calculate the initial rotor flow field to prepare for the subsequent rotor wake and flow field time propagation calculations. Step 4: Update the geometric position and flight state parameters of the rotor blade surface mesh at the current time step, and calculate the rotor wake structure and rotor flow field information at the current time step using the finite difference scheme based on the initial rotor wake and the initial rotor flow field. Step 5: Calculate the blade aerodynamic force based on the rotor flow field information at the current time step, and construct the rotor dynamic equation to solve the aeroelastic response. Determine whether the aeroelastic response converges. If it does not converge, use the rotor wake structure and rotor flow field information at the current time step as the rotor initial wake and rotor initial flow field, and repeat steps 4 to 5 until the blade aeroelastic response converges. Step 6: Calculate the load for the current time step based on the calculated blade aerodynamic force and aeroelastic response; Step 7: Determine the relationship between the load calculation time and the set calculation time. If the load calculation time is less than the set calculation time, repeat steps 4 to 6. If the load calculation time is equal to the set calculation time, the overall rotor wake calculation ends, and the rotor wake structure, rotor flow field, blade aerodynamic force, aeroelastic response and load information of the entire calculation process are output.

[0006] Optionally, the rotor system blade mesh geometry features are generated, specifically: Establish a coordinate system based on the helicopter fuselage, rotor hub, and rotor blades, including the fuselage coordinate system ( Its origin is fixed at the helicopter's center of gravity; the rotor hub coordinate system ( Its origin is fixed at the center of the propeller hub; the undeformed coordinate system of the propeller blades ( Its origin is fixed at the flapping hinge; the blade deformation coordinate system ( Its origin is fixed on the pitch axis; Determine the transformation relationships between the coordinate systems of the helicopter fuselage, rotor hub, and rotor blades; the transformation relationship from the fuselage coordinate system to the rotor hub coordinate system is as follows:

[0007] In the formula, This refers to the rotor shaft tilt angle, with tilt being positive. This represents the rotor shaft tilt angle, with the right side being positive. The transformation relationship from the hub coordinate system to the undeformed blade coordinate system is as follows:

[0008] In the formula, To wave the horn, waving it upwards is the correct direction; This is the azimuth angle, with the direction of rotation being positive. The transformation relationship from the undeformed blade coordinate system to the deformed blade coordinate system is as follows:

[0009] In the formula, Install the blade angle.

[0010] By combining the coordinate system and its transformation relationship, and based on the rotor configuration and geometric parameters, the geometric features of the rotor system blade mesh are generated.

[0011] Optionally, when calculating the rotor wake structure and rotor flow field, an unsteady aerodynamic model is established, specifically including: When establishing an unsteady aerodynamic model, the blades are divided into cosine segments along the span. Each segment, the amount of circulation attached in each segment. For a constant value, the blade is replaced by an attached vortex along the 1 / 4 chord. The attached circulation changes along the span to generate a trailing vortex system that breaks off from the trailing edge of the blade to form a trailing vortex. The control point is located at the intersection of the 3 / 4 chord and the centerline of the blade section. Subsequently, the trailing vortex condenses and rolls up to form a tip vortex with a positive vortex value.

[0012] Optionally, the rotor wake structure and rotor flow field are calculated based on an unsteady aerodynamic model, specifically including: When calculating the rotor wake structure, the fourth-order Runge-Kutta scheme is used to solve the rotor wake structure at the current time step;

[0013] In the formula, This refers to the rotor wake structure at the current time step. This is the initial rotor wake structure; The time step corresponding to the initial rotor wake. For a single calculation step; For time step The rotor wake structure is The velocity field at that time; After solving the rotor wake structure at the current time step using the fourth-order Runge-Kutta scheme, the mutual induced and self-induced velocities in the rotor wake vortex line are calculated using the induced velocity calculation formula. The induced velocity caused by the blade circulation and the free flow of the rotor are added to obtain the rotor flow field at the current time step.

[0014] Optionally, the blades are divided along the spanwise direction into cosine segments. When calculating the spanwise radius of a station segment, the unequal segmentation method is used:

[0015] In the formula, The radius of the starting position for the propeller airfoil. Where is the blade radius, This represents the number of blade segments.

[0016] Optionally, the process of calculating the aerodynamic forces of the blades is as follows: Calculate the circulation of each section of the blade according to the circulation calculation formula. The formula for calculating circulation is:

[0017] In the formula,

[0018]

[0019]

[0020] It is the influence coefficient of the j-th attached vortex on the i-th control point; Let be the influence coefficient of the j-th trailing vortex on the i-th control point; Let be the spanwise width of the blade segment where the i-th control point is located. The lift coefficient, Let z be the z-direction component of the free flow velocity at the i-th control point. Let z be the z-direction component of the induced velocity of the tip vortex at the i-th control point. Let y be the resultant velocity component at the i-th control point. Let be the total velocity of the blade profile corresponding to the i-th control point; The total velocity of the calculated profile is based on blade circulation, rotor wake structure, and rotor flow field. Total velocity of the profile For induced velocity Free flow velocity Rotor blade rotation speed The sum of the three parts, where the induced velocity at any point on the blade is... The calculation formula is:

[0021] The number of rotor blades, Number of segments for the far tail The number of blade spanwise segments. The number of segments for the trailing vortex. Let $\frac{j}{j}$ be the induced velocity of the $j$ segment of the vortex line on the far wake of the $i$-th blade at that point. Let $\frac{i}{j}$ be the induced velocity at that point due to the circulation attached to the $j$ segment of the $i$-th blade. The velocity induced by the trailing vortex of the j-th segment of the i-th blade at this point; the free flow velocity. The rotor blade rotation speed is obtained from the wind speed. It is obtained from the rotor speed and the corresponding profile station position; Based on the total velocity of the profile Calculate the aerodynamic forces of the blades: First, calculate the effective angle of attack of the profile. , Based on the Weissinger-L lifting surface theory, the unsteady lift coefficient, drag coefficient, and pitching moment coefficient of the airfoil profile are calculated, and thus the lift per unit length of the airfoil profile is obtained. ,resistance and pitch moment The aerodynamic force per unit length of the blade deformation coordinate system is obtained through coordinate transformation:

[0022] , , These represent the radial force, chordal force, and vertical force per unit length of the cross-section, respectively. The pitching moment per unit length of the cross section; The sweep angle of the cross-section. After coordinate transformation, the aerodynamic force per unit length in the undeformed blade coordinate system is obtained:

[0023] , , These represent the radial force, chordal force, and vertical force per unit length in the undeformed blade coordinate system, respectively. , , These are the torsional bending moment, flapping bending moment, and oscillation bending moment per unit length in the undeformed blade coordinate system, respectively. The angle of the cross-section is the oscillation angle. The waving angle of the cross section.

[0024] Optionally, when calculating the load for the current time step based on the calculated blade aerodynamic and aeroelastic responses, the blade profile... The load at that point is integrated along the section to the blade tip using the following formula:

[0025] In the formula, , , These are the radial force, chordal force, and vertical force of the cross-section, respectively. , , These are the radial force, chordal force, and vertical force of the cross-section, respectively. , , For arbitrary cross-sections The waving, swinging, and stretching displacement at the location; , , cross section The waving, swinging, and stretching displacement at the location.

[0026] In a second aspect, a transient unsteady rotor aerodynamic load calculation device is provided, which performs the transient unsteady rotor aerodynamic load calculation as described in any of the first aspects, the device comprising: The first processing module is used to acquire rotor system parameters, initialize rotor system parameters, generate rotor blade mesh geometry features, generate initial rotor wake based on rotor blade surface mesh geometry position and flight state parameters at the initial time, calculate initial rotor flow field, update rotor blade surface mesh geometry position and flight state parameters at the current time step, and calculate rotor wake structure and rotor flow field information at the current time step. The second processing module is used to calculate the aerodynamic force of the blade and construct the rotor dynamics equation to solve the aeroelastic response, determine whether the aeroelastic response has converged, and update the rotor wake structure and rotor flow field information until the blade aeroelastic response converges. The third processing module is used to calculate the load at the current time step; determine the relationship between the load calculation time and the set calculation time, until the overall rotor wake calculation is completed, and output the rotor wake structure, rotor flow field, blade aerodynamic force, aeroelastic response and load information of the entire calculation process.

[0027] The beneficial effects of this invention are at least as follows: Using the method of this invention, based on the rotor geometry, wake structure, and flow field of the previous time step, the rotor wake structure, rotor flow field, blade aerodynamic forces, and blade aeroelastic response of the next time step can be directly solved using the finite difference method, thereby calculating the rotor aerodynamic loads under transient unsteady conditions. This method does not require complex CFD calculations, has high computational efficiency, and is less dependent on the model. It is applicable to rotorcraft of different configurations under arbitrary flight states and control input variations in the time domain, expanding the scope of application of methods for calculating rotor unsteady aerodynamic loads. Attached Figure Description

[0028] Figure 1 This is a flowchart of the transient unsteady rotor aerodynamic load calculation method of the present invention; Figure 2 This is a diagram of the rotor wake structure for the transient unsteady rotor aerodynamic load calculation method of the present invention; Figure 3 This is a rotor flow field diagram for the transient unsteady rotor aerodynamic load calculation method of the present invention. Detailed Implementation

[0029] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0030] The features and illustrative embodiments of various aspects of the present invention will now be described in detail. Numerous specific details are set forth in the following detailed description to provide a thorough understanding of the invention. However, it will be apparent to those skilled in the art that the invention may be practiced without requiring some of these specific details. The following description of embodiments is merely intended to provide a better understanding of the invention by illustrating examples of the invention. The invention is by no means limited to any specific setups and methods set forth below, but covers any improvements, substitutions, and modifications to structures, methods, and devices without departing from the spirit of the invention. Well-known structures and techniques are not shown in the drawings and the following description to avoid unnecessarily obscuring the invention.

[0031] It should be noted that, unless otherwise specified, the embodiments of the present invention and the features thereof can be combined with each other, and the various embodiments can be referenced and cited from each other.

[0032] The present invention will now be described in detail with reference to the accompanying drawings and specific embodiments.

[0033] This invention presents a method for calculating transient unsteady rotor aerodynamic loads based on a free wake model. It utilizes the finite difference method to solve for the rotor wake structure and rotor flow field at any given time, thereby obtaining the rotor aerodynamic load results in the time domain. Unlike traditional aerodynamic load calculations, this invention's method no longer relies on pre-set flow field information or is limited to the premise of aeroelastic response convergence after one rotor rotation. It can determine the blade's geometric spatial position based on the converged solution of the blade aeroelastic response from the previous time step, and then, combined with the blade circulation distribution, rotor wake structure, and flow field information from the previous time step, calculate the rotor blade's geometric spatial position, circulation distribution, and wake structure for the next time step. Furthermore, the calculation method provided by this invention is performed in the time domain, adapting to arbitrary changes in helicopter flight conditions and control inputs. This invention improves the accuracy of transient rotor aerodynamic load calculations while effectively enhancing computational efficiency and expanding the application scope of rotor aerodynamic load calculation methods.

[0034] This embodiment provides a method for calculating transient unsteady rotor aerodynamic loads, meeting the calculation requirements for rotor aerodynamic loads at any transient moment, such as with varying incoming flow, varying rotational speed, varying control, and varying attitude. (See also...) Figure 1 Specifically, it includes the following steps: Step 1: First, read in the rotor system parameters to obtain the rotor configuration, geometric parameters, and dynamic parameters, and complete the initialization of the rotor system parameters, such as the dimensionless conversion of the rotor blade physical parameters.

[0035] Step 2: Based on the rotor configuration and geometric parameters, generate the rotor system blade mesh geometric features, and read the rotor blade surface mesh geometric position and flight state parameters at the initial moment to determine the spatial position of the rotor blade mesh nodes and control points in the airframe coordinate system; Step 3: Based on the rotor flight state parameters at the initial moment, generate the initial rotor wake. The initial wake is calculated using a steady inflow model. After obtaining the initial wake structure, the initial flow field of the rotor is calculated to prepare for the subsequent rotor wake and flow field time propagation calculations. Step 4: Update the geometric position and flight state parameters of the rotor blade surface mesh at the current time step, and calculate the rotor wake structure and rotor flow field information at the current time step using the finite difference scheme based on the initial rotor wake and the initial rotor flow field. Step 5: Calculate the blade aerodynamic force based on the rotor flow field information at the current time step, and construct the rotor dynamic equation to solve the aeroelastic response. Determine whether the aeroelastic response converges. If it does not converge, use the rotor wake structure and rotor flow field information at the current time step as the rotor initial wake and rotor initial flow field, and repeat steps 4 to 5 until the blade aeroelastic response converges. Step 6: Calculate the load for the current time step based on the calculated blade aerodynamic force and aeroelastic response; Step 7: Determine the relationship between the load calculation time and the set calculation time. If it is less than the set time, repeat steps 4 to 6. If it is equal to the set time, determine that the overall rotor wake calculation is complete and output the rotor wake structure, rotor flow field, blade aerodynamic force, aeroelastic response and load information of the entire calculation process.

[0036] Specifically, one embodiment of the present invention includes the following: 1) Construct a coordinate system: To concisely describe the rotor flow field, it is necessary to establish a coordinate system based on components such as the helicopter fuselage and rotor, including the fuselage coordinate system ( Its origin is fixed at the helicopter's center of gravity; the rotor hub coordinate system ( Its origin is fixed at the center of the propeller hub; the undeformed coordinate system of the propeller blades ( Its origin is fixed at the flapping hinge; the blade deformation coordinate system ( Its origin is fixed on the pitch axis.

[0037] Transformation from body coordinate system to rotor hub coordinate system ( )for:

[0038] This refers to the rotor shaft tilt angle, with tilt being positive. This represents the rotor shaft tilt angle, with the right side being positive.

[0039] From hub coordinate system to blade undeformed coordinate system ( )for:

[0040] To wave the horn, waving it upwards is the correct direction; It is the azimuth angle, with the direction of rotation being positive.

[0041] From blade coordinate system to blade profile coordinate system ( )for:

[0042] Install the blade angle.

[0043] 2) Unsteady aerodynamic model The unsteady aerodynamic model of the blades employs the second-order lift line theory, namely the Weissinger-L lift surface theory. The blades are divided into cosine sections along the spanwise axis. Segment, the amount of circulation attached in each segment As a constant, the blade is replaced by an attached vortex along the 1 / 4 chord. The attached circulation varies along the spanwise, generating a trailing vortex system that detaches from the trailing edge of the blade (i.e., the difference in circulation between two adjacent blade segments), forming a trailing vortex with a trailing vortex angle of 30°. The control point is located at the intersection of the 3 / 4 chord and the blade segment centerline. To more accurately describe the nonlinear changes in blade circulation and load, the blade segments can be divided using an unequal segmentation method, with the segment spanwise station radius as follows:

[0044] The radius of the starting position for the propeller airfoil. Where is the blade radius, This represents the number of blade segments. A schematic diagram of the Weissinger-L lifting surface model in the unsteady aerodynamic model can be found here. Figure 1 .

[0045] 3) Blade circulation calculation The lift per unit length of each blade segment was calculated using the Kuta-Jukovsky theorem.

[0046]

[0047] air density, The total velocity of the cross section.

[0048] At the same time, the lift per unit length of each blade It is also the effective angle of attack of the blade profile. function

[0049] c This refers to the spanwise width of the blade section.

[0050] The effective angle of attack of the blade profile is as follows:

[0051] Let z be the z-direction velocity component of the free flow at the blade control point. , , These are the z-direction velocity components of the induced velocities of the tip vortex, attached vortex, and trailing vortex at the blade control point, respectively. The y-direction component is used to control the net velocity of the control point.

[0052] Based on the induced velocities of the control point by the blade-attached vortex and the trailing vortex, we have:

[0053] The circulation of each section of the blade is calculated as follows:

[0054] In the formula,

[0055]

[0056]

[0057] Based on the blade circulation obtained from the solution, the aerodynamic distribution of the blade is then obtained by solving the rotor wake structure and flow field information.

[0058] 4) Solving the rotor wake structure The equation of motion for the rotor wake can be written as:

[0059] Due to the velocity field It is highly nonlinear, so the equations of motion must be discretized using a finite difference scheme before they can be solved numerically through iteration. When calculating the rotor wake structure, a fourth-order Runge-Kutta scheme is used to solve the rotor wake structure at the current time step; see [link to relevant documentation]. Figure 2 ,

[0060] This refers to the rotor wake structure at the current time step. This is the initial rotor wake structure; The time step corresponding to the initial rotor wake. For a single calculation step; For time step The rotor wake structure is The velocity field at time.

[0061] After solving the rotor wake structure at the current time step using the fourth-order Runge-Kutta scheme, the mutually induced and self-induced velocities in the rotor wake vortex lines can be calculated using the induced velocity calculation formula. Adding the induced velocity caused by the blade circulation and the free flow of the rotor, the rotor flow field at the current time step is obtained. (See [link to relevant documentation]). Figure 3 .

[0062] 5) Solving for the induced velocity The position vector of the separation point between the trailing vortex and the propeller blade in the body coordinate system is: ,in For the time when the wake separates from the blade, it moves freely at the local velocity. For a point where the trailing vortex separates from the blade, the governing equation can be written as:

[0063] For the incoming flow velocity, Let be the induced velocity at that point. Determined by the following formula: .

[0064] The key points of this invention are as follows: 1) To address the need for analyzing transient unsteady rotor aerodynamic loads, this paper addresses the convergence constraint requirement of the aeroelastic response after one revolution of a conventional steady-state rotor. Based on the original wake motion equations, and with time as the main variable, the paper employs the finite difference method to solve the rotor wake equations in a time-series manner. This solves the problem that conventional steady-state rotor wake models are not applicable to rotor speed variations and transient requirements.

[0065] 2) For rotor transient processes, any rotor motion time history can be set, realizing rotor aerodynamic analysis of transient processes such as variable incoming flow, variable speed, variable control, and variable attitude.

[0066] The above description merely illustrates embodiments of the present invention and is quite specific and detailed; however, it should not be construed as limiting the scope of the patent. It should be noted that those skilled in the art can make various modifications and improvements without departing from the concept of the present invention, and these all fall within the scope of protection of the present invention. Furthermore, any parts of the present invention not described in detail are conventional techniques.

Claims

1. A method for calculating transient unsteady rotor aerodynamic loads, characterized in that, The method includes: Obtain rotor system parameters and initialize them. Generate rotor blade mesh geometry features. Based on the rotor blade surface mesh geometry position and flight state parameters at the initial time, generate the rotor initial wake and calculate the rotor initial flow field. Update the rotor blade surface mesh geometry position and flight state parameters at the current time step, and calculate the rotor wake structure and rotor flow field information at the current time step. Calculate the aerodynamic force of the blade and construct the rotor dynamics equation to solve the aeroelastic response. Determine whether the aeroelastic response has converged, and update the rotor wake structure and rotor flow field information until the blade aeroelastic response converges. Calculate the load at the current time step; determine the relationship between the load calculation time and the set calculation time, until the overall rotor wake calculation is completed, and output the rotor wake structure, rotor flow field, blade aerodynamic force, aeroelastic response and load information of the entire calculation process.

2. The method according to claim 1, characterized in that, The method specifically includes: Step 1: Obtain the rotor system parameters and complete the initialization of the rotor system parameters, including rotor configuration, geometric parameters, and dynamic parameters; Step 2: Based on the rotor configuration and geometric parameters, generate the rotor system blade mesh geometric features, and read the rotor blade surface mesh geometric position and flight state parameters at the initial moment; Step 3: Based on the initial geometric position of the rotor blade surface grid and flight state parameters, generate the initial rotor wake and calculate the initial rotor flow field to prepare for the subsequent rotor wake and flow field time propagation calculations. Step 4: Update the geometric position and flight state parameters of the rotor blade surface mesh at the current time step, and calculate the rotor wake structure and rotor flow field information at the current time step using the finite difference scheme based on the initial rotor wake and the initial rotor flow field. Step 5: Calculate the blade aerodynamic force based on the rotor flow field information at the current time step, and construct the rotor dynamic equation to solve the aeroelastic response. Determine whether the aeroelastic response converges. If it does not converge, use the rotor wake structure and rotor flow field information at the current time step as the rotor initial wake and rotor initial flow field, and repeat steps 4 to 5 until the blade aeroelastic response converges. Step 6: Calculate the load for the current time step based on the calculated blade aerodynamic force and aeroelastic response; Step 7: Determine the relationship between the load calculation time and the set calculation time. If the load calculation time is less than the set calculation time, repeat steps 4 to 6. If the load calculation time is equal to the set calculation time, the overall rotor wake calculation ends, and the rotor wake structure, rotor flow field, blade aerodynamic force, aeroelastic response and load information of the entire calculation process are output.

3. The method according to claim 2, characterized in that, Generate the rotor system blade mesh geometry features, specifically: Establish a coordinate system based on the helicopter fuselage, rotor hub, and rotor blades, including the fuselage coordinate system ( Its origin is fixed at the helicopter's center of gravity; the rotor hub coordinate system ( Its origin is fixed at the center of the propeller hub; Undeformed blade coordinate system ( Its origin is fixed at the flapping hinge; the blade deformation coordinate system ( Its origin is fixed on the pitch axis; Determine the transformation relationships between the coordinate systems of the helicopter fuselage, rotor hub, and rotor blades; the transformation relationship from the fuselage coordinate system to the rotor hub coordinate system is as follows: In the formula, This refers to the rotor shaft tilt angle, with tilt being positive. This represents the rotor shaft tilt angle, with the right side being positive. The transformation relationship from the hub coordinate system to the undeformed blade coordinate system is as follows: In the formula, To wave the horn, waving it upwards is the correct direction; This is the azimuth angle, with the direction of rotation being positive. The transformation relationship from the undeformed blade coordinate system to the deformed blade coordinate system is as follows: In the formula, Install the blade angle. By combining the coordinate system and its transformation relationship, and based on the rotor configuration and geometric parameters, the geometric features of the rotor system blade mesh are generated.

4. The method according to claim 2, characterized in that, When calculating the rotor wake structure and rotor flow field, an unsteady aerodynamic model is established, which specifically includes: When establishing an unsteady aerodynamic model, the blades are divided into cosine segments along the spanwise direction. Each segment, the amount of circulation attached in each segment. For a constant value, the blade is replaced by an attached vortex along the 1 / 4 chord. The attached circulation changes along the span to generate a trailing vortex system that breaks off from the trailing edge of the blade to form a trailing vortex. The control point is located at the intersection of the 3 / 4 chord and the centerline of the blade section. Subsequently, the trailing vortex condenses and rolls up to form a tip vortex with a positive vortex value.

5. The method according to claim 4, characterized in that, The rotor wake structure and rotor flow field were calculated based on an unsteady aerodynamic model, specifically including: When calculating the rotor wake structure, the fourth-order Runge-Kutta scheme is used to solve the rotor wake structure at the current time step; In the formula, This refers to the rotor wake structure at the current time step. This is the initial rotor wake structure; The time step corresponding to the initial rotor wake. For a single calculation step; For time step The rotor wake structure is The velocity field at that time; After solving the rotor wake structure at the current time step using the fourth-order Runge-Kutta scheme, the mutual induced and self-induced velocities in the rotor wake vortex line are calculated using the induced velocity calculation formula. The induced velocity caused by the blade circulation and the free flow of the rotor are added to obtain the rotor flow field at the current time step.

6. The method according to claim 4, characterized in that, Divide the blade along its span according to cosine. When calculating the spanwise radius of a station segment, the unequal segmentation method is used: In the formula, The radius of the starting position for the propeller airfoil. Where is the blade radius, This represents the number of blade segments.

7. The method according to claim 1, characterized in that, The process of calculating the aerodynamic forces of the propeller blades is as follows: Calculate the circulation of each section of the blade according to the circulation calculation formula. The formula for calculating circulation is: In the formula, It is the influence coefficient of the j-th attached vortex on the i-th control point; Let be the influence coefficient of the j-th trailing vortex on the i-th control point; Let be the spanwise width of the blade segment where the i-th control point is located. The lift coefficient, Let z be the z-direction component of the free flow velocity at the i-th control point. Let z be the z-direction component of the induced velocity of the tip vortex at the i-th control point. Let y be the resultant velocity component at the i-th control point. Let be the total velocity of the blade profile corresponding to the i-th control point; The total velocity of the calculated profile is based on blade circulation, rotor wake structure, and rotor flow field. Total velocity of the profile For induced velocity Free flow velocity Rotor blade rotation speed The sum of the three parts, where the induced velocity at any point on the blade is... The calculation formula is: The number of rotor blades, Number of segments for the far tail The number of blade spanwise segments. The number of segments for the trailing vortex. Let $\frac{j}{j}$ be the induced velocity of the $j$ segment of the vortex line on the far wake of the $i$-th blade at that point. Let be the induced velocity at that point caused by the circulation attached to the j-th segment of the i-th blade. The velocity induced by the trailing vortex of the j-th segment of the i-th blade at this point; the free flow velocity. The rotor blade rotation speed is obtained from the wind speed. It is obtained from the rotor speed and the corresponding profile station position; Based on the total velocity of the profile Calculate the aerodynamic forces of the blades: First, calculate the effective angle of attack of the profile. , Based on the Weissinger-L lifting surface theory, the unsteady lift coefficient, drag coefficient, and pitching moment coefficient of the airfoil profile are calculated, and thus the lift per unit length of the airfoil profile is obtained. ,resistance and pitch moment The aerodynamic force per unit length of the blade deformation coordinate system is obtained through coordinate transformation: , , These represent the radial force, chordal force, and vertical force per unit length of the cross-section, respectively. The pitching moment per unit length of the cross section; The sweep angle of the cross-section. After coordinate transformation, the aerodynamic force per unit length in the undeformed blade coordinate system is obtained: , , These represent the radial force, chordal force, and vertical force per unit length in the undeformed blade coordinate system, respectively. , , These are the torsional bending moment, flapping bending moment, and oscillation bending moment per unit length in the undeformed blade coordinate system, respectively. The angle of the cross-section is the oscillation angle. The waving angle of the cross section.

8. The method according to claim 2, characterized in that, When calculating the load for the current time step based on the calculated blade aerodynamic forces and aeroelastic response, the blade profile is considered. The load at that point is integrated along the section to the blade tip using the following formula: In the formula, , , These are the radial force, chordal force, and vertical force of the cross-section, respectively. , , These are the radial force, chordal force, and vertical force of the cross-section, respectively. , , For arbitrary cross-sections The waving, swinging, and stretching displacement at the location; , , cross section The waving, swinging, and stretching displacement at the location.

9. A device for calculating transient unsteady rotor aerodynamic loads, characterized in that, The apparatus for performing transient unsteady rotor aerodynamic load calculations according to any one of claims 1 to 8 includes: The first processing module is used to acquire rotor system parameters, initialize rotor system parameters, generate rotor blade mesh geometry features, generate initial rotor wake based on rotor blade surface mesh geometry position and flight state parameters at the initial time, calculate initial rotor flow field, update rotor blade surface mesh geometry position and flight state parameters at the current time step, and calculate rotor wake structure and rotor flow field information at the current time step. The second processing module is used to calculate the aerodynamic force of the blade and construct the rotor dynamics equation to solve the aeroelastic response, determine whether the aeroelastic response has converged, and update the rotor wake structure and rotor flow field information until the blade aeroelastic response converges. The third processing module is used to calculate the load at the current time step; determine the relationship between the load calculation time and the set calculation time, until the overall rotor wake calculation is completed, and output the rotor wake structure, rotor flow field, blade aerodynamic force, aeroelastic response and load information of the entire calculation process.

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