Rotor craft flow field numerical simulation method, equipment, medium and product

Abstract

The invention discloses a rotor craft flow field numerical simulation method and device, a medium and a product, and relates to the technical field of aircraft flow field data processing, and the method comprises the steps: building a rotor craft model based on a rotor craft under a given flight condition; generating a hybrid grid system based on the rotorcraft model; according to the hybrid grid system, performing flow field CFD numerical simulation on the rotorcraft to obtain conservation variable information of a blade area; according to the conservation variable information of the blade area, an AASM method is adopted to calculate and obtain an unsteady momentum source item of the blade area; and according to the unsteady momentum source item of the blade area, rotor aerodynamic characteristic data and flow field data under the given flight working condition are determined. By coupling the AASM method and the CFD simulation, the precision and efficiency of calculating the aerodynamic characteristic data and the flow field data of the rotor wing can be improved, and the requirements of carrying out complex full-aircraft interference with great resource consumption and maneuvering flight state research are met.

Description

Technical Field

[0001] The present application relates to the technical field of aircraft flow field data processing, and in particular to a method, device, medium and product for numerical simulation of rotorcraft flow field. Background Art

[0002] As the most important component of a rotorcraft, the rotor provides most of the lift and control force. High-precision calculation of the rotor flow field plays an important role in the study of aircraft performance, rotor aerodynamic loads, flight dynamics, and aeroacoustics. However, accurate simulation of the aerodynamic interference, blade-vortex interference, and maneuvering calculations generated by the rotor under different flight conditions often requires a large computer cluster and consumes a lot of computing resources. Therefore, it is of great significance to accurately and efficiently simulate the unsteady flow field of the rotor.

[0003] At present, the calculation of rotor flow field can be basically divided into three categories. The first is the wake method based on vortex theory. This method has a small amount of model calculation and relies heavily on empirical parameters, which is difficult to meet the increasingly complex engineering needs. The second is the full-blade analytical method based on motion nested grids. This method uses volume grids to geometrically model the real blade model. The simulation accuracy is high, but it consumes a lot of computing resources, especially for flow field simulations that require small time steps and large grid density, such as blade-vortex interference and full-machine maneuvers. The third is the virtual disk model method based on the steady momentum source method. This method uses a thin body grid disk to replace the rotor. There is no need to generate blade body grids, which reduces the number of grids in the computational domain. However, due to the time averaging of the periodic effect of the blade on the surrounding air, the unsteady load characteristics of the rotor cannot be effectively captured. In order to make the virtual disk model suitable for the simulation of the unsteady flow field of the rotor, there is an unsteady momentum source method in the related technology. This method no longer regards the rotor as a whole actuator disk, but only adds source terms to the area where the blade is located in each time step. The actuating line method and the actuating surface method are developed in turn. According to the different distribution methods of source items, it is mainly divided into two categories: the actuation line method and the actuation surface method. Among them, the actuation line method distributes the source items on a line that simulates the blade. Since the grids selected for the previous and next time steps do not overlap, the flow field continuity is poor and there are large disturbances. The actuation surface method uses empirical functions or Gaussian functions to distribute the load along the chord direction of the blade to prevent singular values ​​from appearing at the leading and trailing edges of the blade. Related technologies also combine the actuation surface method with a dynamic inflow model, and couple the actuation surface method with a predetermined / free wake model. Most of them are applied in the field of wind turbines. In recent years, an unsteady momentum source method based on three-dimensional equidistant Gaussian kernel has been developed. This method distributes source terms in the span, chord and chord-normal directions of the blade, and uses a smearing parameter to control the interaction between the units. This method was then combined with Star CCM (flow field simulation software), and the applicability of the Sikorsky S-76 (Sikorsky "S-76" helicopter) rotor hovering state was evaluated using a cut volume grid. The results show that the prediction of the blade tip load under the action of blade-vortex interference is overestimated.

[0004] At present, the research on the actuation line and actuation surface method is still limited to the research on wind turbines. In the related technology, the actuation line model and the three-dimensional panel method model are combined to develop an improved actuation surface model, which is applied to the study of the aerodynamic characteristics of wind turbines. This model is based on the real geometric shape of the blade, and uses the three-dimensional panel method to directly construct the point source and dipole distribution on the blade surface. The calculation accuracy is good in the near-wake area of ​​the blade. Compared with wind turbines, the three-dimensional effect of the blade caused by the tip vortex of the rotorcraft is more significant. The blades need to be flapped and pitch-changed during flight. High-performance blades even have nonlinear negative torsion and chord length, and have very complex geometric shapes. The traditional momentum source method is difficult to meet the calculation of the rotorcraft flow field. However, to date, the development of unsteady momentum source methods for rotorcraft has not yet been carried out, and the simulation is limited to the use of time-averaged or time-accurate virtual disk methods to simulate the rotor flow field. Faced with complex flight conditions, the rotor aerodynamic load and other rotor aerodynamic characteristic data calculated by this method often have large fluctuations, low accuracy, and low efficiency, and cannot meet the needs of conducting complex full-aircraft interference and maneuvering flight state research that consumes a lot of resources.

[0005] In summary, how to provide a high-precision and high-efficiency numerical simulation method for rotorcraft flow field to improve the accuracy and efficiency of calculating rotor aerodynamic characteristics data and flow field data has become a technical problem that needs to be urgently solved in this field. Summary of the invention

[0006] The purpose of this application is to provide a method, equipment, medium and product for numerical simulation of the flow field of a rotorcraft, which can improve the accuracy and efficiency of calculating the aerodynamic characteristics data and flow field data of the rotor, and meet the needs of conducting research on complex full-machine interference and maneuverable flight states with extremely high flight resource consumption.

[0007] To achieve the above objectives, this application provides the following solutions:

[0008] In a first aspect, the present application provides a method for numerically simulating a flow field of a rotorcraft, the method comprising:

[0009] Based on the rotorcraft under given flight conditions, a rotorcraft model is established; the rotorcraft model is a three-dimensional CAD model including an aircraft body model and a rotor blade model.

[0010] Based on the rotorcraft model, a hybrid grid system is generated; the hybrid grid system includes a grid structure corresponding to the aircraft body model and a grid structure corresponding to the rotor disc model, and the hybrid grid system is used to provide the grid structure and boundary conditions required for flow field calculation.

[0011] According to the hybrid grid system, the flow field CFD numerical simulation of the rotorcraft is performed to obtain the conservation variable information of the blade area; the conservation variable information includes the velocity field information and the density field information of the blade area.

[0012] According to the conserved variable information of the blade region, the unsteady momentum source term of the blade region is calculated by using the AASM method.

[0013] According to the unsteady momentum source term of the blade area, the aerodynamic characteristic data and flow field data of the rotor under a given flight condition are determined.

[0014] In a second aspect, the present application provides a computer device, comprising: a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor executes the computer program to implement any of the above-described methods for numerically simulating the flow field of a rotorcraft.

[0015] In a third aspect, the present application provides a computer-readable storage medium having a computer program stored thereon, which, when executed by a processor, implements any of the above-described methods for numerically simulating the flow field of a rotorcraft.

[0016] In a fourth aspect, the present application provides a computer program product, including a computer program, which, when executed by a processor, implements any of the above-mentioned methods for numerical simulation of the rotorcraft flow field.

[0017] According to the specific embodiments provided in this application, this application discloses the following technical effects:

[0018] The present application provides a method, device, medium and product for numerical simulation of the flow field of a rotorcraft, which obtains the conservation variable information of the blade area by performing CFD numerical simulation of the flow field of the rotorcraft, and uses the AASM method to calculate the unsteady momentum source term of the blade area based on the conservation variable information of the blade area, thereby determining the rotor aerodynamic characteristic data and flow field data of the rotorcraft. It can be seen that the present application couples the AASM method with CFD simulation and applies it to the scenario of high-fidelity flow field calculation of rotorcraft, which can effectively improve the accuracy and efficiency of calculating the rotor aerodynamic characteristic data and flow field data, and meet the needs of conducting complex full-machine interference and maneuvering flight state research with extremely high resource consumption. BRIEF DESCRIPTION OF THE DRAWINGS

[0019] In order to more clearly illustrate the embodiments of the present application or the technical solutions in the prior art, the drawings required for use in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present application. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0020] Figure 1 A schematic flow chart of a method for numerically simulating a rotorcraft flow field provided in accordance with an embodiment of the present application.

[0021] Figure 2 A schematic diagram of the process of AASM and CFD coupling calculation provided in one embodiment of the present application.

[0022] Figure 3 A schematic diagram of an AASM method provided in accordance with an embodiment of the present application.

[0023] Figure 4 A schematic diagram of the conversion relationship between the global coordinate system and the propeller disc coordinate system provided in an embodiment of the present application.

[0024] Figure 5 A schematic diagram of the conversion relationship between the propeller disc coordinate system and the propeller blade coordinate system provided in one embodiment of the present application.

[0025] Figure 6 A schematic diagram of the aerodynamic force and incoming flow of a blade section provided in one embodiment of the present application.

[0026] Figure 7 A schematic diagram of the variation of the tension coefficient with the azimuth angle during the AASM method, the traditional unsteady momentum source method and the advanced actuating surface method tests provided in one embodiment of the present application.

[0027] Figure 8 A schematic diagram of the variation of lift with azimuth angle when r / R=0.50 for the AASM method, the traditional unsteady momentum source method and the motion nested grid method provided in one embodiment of the present application.

[0028] Fig. 9 A schematic diagram of the variation of lift with azimuth angle when r / R=0.65 for the AASM method, the traditional unsteady momentum source method and the motion nested grid method provided in one embodiment of the present application.

[0029] Fig.10 A schematic diagram of the variation of lift with azimuth angle when r / R=0.89 for the AASM method, the traditional unsteady momentum source method and the motion nested grid method provided in one embodiment of the present application.

[0030] Fig.11A schematic diagram of the variation of lift with azimuth angle when r / R=0.99 for the AASM method, the traditional unsteady momentum source method and the motion nested grid method provided in one embodiment of the present application. DETAILED DESCRIPTION

[0031] The following will be combined with the drawings in the embodiments of the present application to clearly and completely describe the technical solutions in the embodiments of the present application. Obviously, the described embodiments are only part of the embodiments of the present application, not all of the embodiments. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of this application.

[0032] In order to make the above-mentioned objects, features and advantages of the present application more obvious and easy to understand, the present application is further described in detail below in conjunction with the accompanying drawings and specific implementation methods.

[0033] like Figure 1 As shown, this embodiment provides a method for numerically simulating the flow field of a rotorcraft, and the method for numerically simulating the flow field of a rotorcraft specifically includes the following steps:

[0034] Step S1: Based on the rotorcraft under given flight conditions, a rotorcraft model is established, wherein the rotorcraft model is a three-dimensional CAD (Computer Aided Design) model corresponding to the rotorcraft, including an aircraft body model and a rotor blade model.

[0035] Step S2: Generate a hybrid grid system based on the rotorcraft model. The hybrid grid system includes a grid structure corresponding to the aircraft body model and a grid structure corresponding to the rotor blade model. The hybrid grid system is used to provide the grid structure and boundary conditions required for flow field calculation.

[0036] Step S3: Perform CFD (Computational Fluid Dynamics) numerical simulation of the flow field of the rotorcraft according to the hybrid grid system to obtain the conserved variable information of the blade area, wherein the conserved variable information includes the velocity field information and the density field information of the blade area.

[0037] Step S4: According to the conserved variable information of the blade area, the unsteady momentum source term of the blade area is calculated by using the Advanced Actuator Surface Method (AASM) method.

[0038] Step S5: determining the rotor aerodynamic characteristic data and flow field data under given flight conditions according to the unsteady momentum source term of the blade area.

[0039] In this embodiment, step S1 establishes a rotorcraft model based on the rotorcraft under given flight conditions, and specifically includes the following steps:

[0040] For the rotorcraft under given flight conditions, according to the basic dimensions of the rotorcraft, CAD software is used to respectively establish the aircraft body model and the rotor disc model corresponding to the rotorcraft to obtain the rotorcraft model.

[0041] In this embodiment, step S2 generates a hybrid grid system based on the rotorcraft model, which specifically includes the following steps:

[0042] Step S21, based on the aircraft body model and the background grid area, an array advancement method is used to generate an unstructured tetrahedral grid block; the unstructured tetrahedral grid block is a grid structure corresponding to the aircraft body model, and the unstructured tetrahedral grid block includes a plurality of tetrahedral grid units.

[0043] Step S22: Based on the rotor disk model and the rotor disk area, a sweeping method is used to generate a hexahedral grid block; the hexahedral grid block is a grid structure corresponding to the rotor disk model, and the hexahedral grid block includes a plurality of hexahedral grid units.

[0044] Step S23, the surface mesh units corresponding to the external flow field and the surface mesh units corresponding to the disk flow field are shared topology at the disk interface, and pyramid mesh units are used to smoothly transition the tetrahedral mesh units and the hexahedral mesh units; wherein the tetrahedral mesh units, the hexahedral mesh units, the surface mesh units and the pyramid mesh units are different mesh types in the hybrid mesh system.

[0045] In this embodiment, step S3 performs flow field CFD numerical simulation on the rotorcraft according to the hybrid grid system to obtain conservation variable information of the blade area, which specifically includes the following steps:

[0046] Based on the given flight conditions and incoming flow conditions, the flow field CFD numerical simulation of the rotorcraft is performed according to the hybrid grid system, and the U-RANS (Unsteady Reynolds-Averaged Navier-Stokes equations) momentum equation is used as the main control equation. The U-RANS momentum equation is solved by time advancement to obtain the velocity field information and density field information of the blade area in each iteration step.

[0047] In this embodiment, step S4 calculates the unsteady momentum source term of the blade region using the AASM method according to the conserved variable information of the blade region, and specifically includes the following steps:

[0048] Step S41, performing three-dimensional integral sampling on the velocity field information and density field information of the blade area in each iteration step to obtain a three-dimensional integral sampling result.

[0049] Step S42: Using the three-dimensional integral sampling result as the input of the AASM method, constructing a velocity triangle on each blade section in the blade area; the velocity triangle is used to determine the blade element force of the blade section.

[0050] Step S43, according to the velocity triangle on each blade section, based on the Blade Element Theory (BET), calculate the blade element force of each blade section to obtain a blade element force calculation result.

[0051] Step S44: convert the blade element force calculation result of each blade section into a momentum source term, and convolve them with anisotropic three-dimensional Gaussian kernels to obtain an unsteady momentum source term of the blade region.

[0052] In this embodiment, step S5 determines the rotor aerodynamic characteristic data and flow field data under a given flight condition according to the unsteady momentum source term of the blade region, and specifically includes the following steps:

[0053] The unsteady momentum source term of the blade region is added to the residual calculation of the U-RANS momentum equation for iterative calculation, and the rotor aerodynamic characteristic data and the flow field data are output in real time.

[0054] According to the task requirements of calculating and analyzing the aerodynamic characteristics of rotorcraft, the present invention couples the AASM method with the CFD simulation method based on a hybrid grid system, and applies it to the calculation of high-fidelity rotor flow fields. Among them, a hybrid grid generation strategy based on the sweeping method is used to create a rotor disc grid to be equivalent to the real rotor blades. A rotor blade aerodynamic calculation model is established based on blade element theory and anisotropic three-dimensional Gaussian kernel. A rotor flow field numerical simulation method is established based on the U-RANS momentum equation, and coupled with the AASM method, a study on the flow field characteristics, interference effects and aerodynamic response of maneuvering subjects of rotorcraft is carried out. Specifically including the following contents:

[0055] (1) Modeling of the aircraft body and rotor disc. First, a thin cylindrical rotor disc model and an aircraft body model are established to provide the rotor momentum source application area and the fuselage solid model respectively.

[0056] (2) Generation of hybrid grid system for flow field in computational domain. For the aircraft body model and background grid area, the array propulsion method is used to generate unstructured tetrahedral grid units; for the rotor disc model and rotor disc area, the sweeping method is used to generate hexahedral grid units; the surface grid units of the external flow field and the disc flow field share topology at the disc interface, and pyramid grid units are used for transition to form a hybrid grid system, which provides grids and boundary conditions for CFD numerical simulation calculations.

[0057] (3) CFD numerical simulation of rotorcraft flow field. For the specified incoming flow conditions and flight conditions, the generated hybrid grid system and the given boundary conditions are used to carry out CFD numerical simulation of rotorcraft flow field. By solving the time-marching U-RANS momentum equation, the conservation variable information of the blade area at each iteration step is obtained.

[0058] (4) Calculation of the unsteady momentum source term in the blade area. In each iteration step, the velocity field information and density field information extracted from the CFD solution are sampled in three dimensions as the input of the AASM method. The velocity triangle is constructed in each blade section, and the blade element force of the blade section is calculated according to the blade element theory. After being converted into a momentum source term, it is convolved with the anisotropic three-dimensional Gaussian kernel and added to the source term in the U-RANS momentum equation.

[0059] (5) Coupling simulation and aerodynamic characteristics analysis of the rotor flow field. After bidirectional coupling iterative calculations of the CFD and AASM methods, the unsteady flow field of the rotor forms a periodic oscillation convergence trend, and the aerodynamic characteristics and flow field data of the rotor under given flight conditions are obtained through the post-processing interface. Among them, the post-processing interface is an interface set for post-processing. Post-processing refers to the real-time output of the aerodynamic characteristics of the rotor, single blade, and single blade section during the calculation process, including aerodynamic loads, induced speeds, aerodynamic coefficients, blade-vortex interference positions, etc. If there is a fuselage, data such as pressure on the fuselage surface and specified surface monitoring points can also be output. After the calculation is completed, all conservation variable information of the computational domain flow field is output.

[0060] In order to make the technical solution of this embodiment clearer, the specific implementation process of this embodiment is described below in the form of examples.

[0061] (1) First, a hybrid grid system suitable for AASM calculation is generated. Taking the flow field of the rotor-fuselage combination as an example, the following steps are specifically included:

[0062] A1. Generate unstructured tetrahedral mesh blocks based on the aircraft body model. First, create the far-field boundary, remove the propeller disk part and the fuselage part through subtraction operation to form an external calculation domain. The upper and lower boundaries of the calculation domain are 4R and 10R away from the rotor center, respectively, and the left and right boundaries are 5R away from the rotor center. The front and rear boundaries in the incoming flow direction are 5R and 10R away from the rotor center, respectively. Then, the far-field boundary and the fuselage surface boundary in the discretized calculation domain are generated to generate the surface mesh, and then the array advancement method is used to generate the unstructured tetrahedral unit, and the body-fitting mesh is generated for the boundary layer flow area of ​​the fuselage surface.

[0063] A2. Generate a hexahedral mesh block based on the rotor disc model. First, a thin disc model is established according to the rotor radius and root cutting position. The upper and lower surfaces of the disc are set as source and receiver items respectively, and the disc is swept in the thickness direction. The disc is cut into 5 equally spaced grids. 360 nodes are distributed along the circumference of the blade and 2n nodes are distributed along the span of the blade by surface mapping between the root loop and the tip loop. n is the number of blade element micro-segments of the blade. It is necessary to satisfy n>20 to retain the blade span geometry information as much as possible, and avoid expanding the influence range of the blade element when performing the convolution of the blade element force and the Gaussian kernel, resulting in the loss of local aerodynamic characteristics.

[0064] A3. Shared topology and mesh encryption at the interface. In order to ensure the continuity of the flow field in the computational domain, the disk grid blocks and the background grid blocks need to reconstruct the geometric topology at the grid boundaries and share the grid nodes at the interface. Since the hexahedral grid cells and the tetrahedral grid cells cannot be directly connected, the hexahedral grid cells and the tetrahedral grid cells in this embodiment use pyramid grid cells for smooth transition. Mesh encryption is performed in the area 0.2R above and below the disk and 0.2R extending radially to better capture the near-wake characteristics of the rotor. Similarly, mesh encryption is performed in the rotor wake area and near the fuselage.

[0065] A4. Given the boundary conditions required for CFD calculation, for the external compressible flow boundary far enough away from the rotorcraft, a non-reflection treatment is adopted, that is, the disturbance wave will not be reflected back into the flow field. Given the pressure far-field boundary conditions, based on the extrapolation of the Riemann invariant, it is assumed that there is an irrotational flow in the boundary normal. The far-field boundary normal velocity V n The amplitude of is calculated according to the characteristic line theory:

[0066]

[0067] R + and R - are positive and negative Riemann invariants, respectively, and are defined as follows:

[0068]

[0069] Among them, R +It is obtained by interpolating the internal field grid cells along the outflow characteristic line, R - It is calculated using free flow along the inflow characteristic line, γ is the specific heat ratio, n represents the normal direction, ext represents the value of the adjacent grid unit at the inner field boundary, ∞ represents the free flow value, and c is the local sound speed, which is calculated as follows:

[0070]

[0071] According to the normal velocity V n and the corresponding Mach number M n∞ The entropy s and tangential velocity V on the boundary can be obtained from the flow value or the internal field value. t , there are four specific situations:

[0072] 1) Subsonic inflow boundary conditions (M n∞ <1,V n <0), then we have the following formula:

[0073]

[0074] 2) Subsonic outflow boundary conditions (M n∞ <1,V n >0), then we have the following formula:

[0075]

[0076] 3) Supersonic inflow boundary conditions (M n∞ >1,V n <0), then we have the following formula:

[0077]

[0078] 4) Supersonic outflow boundary conditions (M n∞ >1,V n >0), then we have the following formula:

[0079]

[0080] According to the V n ,c,s,V t Values ​​and Relationships and The values ​​of u, v, p, and ρ on the boundary can be obtained, and the values ​​of each conserved variable on the boundary can be further calculated.

[0081] Given the physical surface boundary conditions on the fuselage surface, under the viscous flow state, the relative velocity of the airflow on the physical surface is zero, that is, V t = 0. In addition, the thermodynamic and kinetic boundary conditions on the surface use adiabatic conditions and no penetration conditions respectively, and the normal gradients of temperature and pressure are zero, that is, and

[0082] The internal surface boundary conditions given on the propeller-disk interface have no effect during the calculation, that is, the propeller-disk area and other areas are regarded as the same calculation domain.

[0083] (2) Input CFD and AASM configuration files and mesh files, and carry out CFD numerical simulation to obtain the velocity field and density field information of the blade area, such as Figure 2 As shown, the specific steps include:

[0084] B1. Read the CFD configuration file. The configuration file is organized in JSON format, describing the grid block and its file name, flow field initial parameters (reference value, incoming flow conditions, etc.), AASM coupling information (propeller disk index, linked propeller disk configuration file name, maneuvering task input, etc.), solution model parameters (space / time discrete format, turbulence model, CFL number, etc.), solution control parameters (maximum time step, maximum pseudo-iteration step, time step, residual convergence standard, etc.), post-processing and output settings.

[0085] B2. Read the AASM configuration file. The AASM configuration consists of the blade profile parameter file (twist table, chord table, airfoil C81 table) and the rotor overall parameter file. The torsion table and chord table contain geometric information such as the torsion angle and local chord length of each section position along the span of the blade. The airfoil C81 table contains the lift / drag coefficient of each airfoil at different Mach numbers and incoming flow angles of attack. The rotor overall parameter file provides rotor speed, rotation direction, radius, root cut, disc thickness, disc direction, disc center global coordinates, number of blades, cyclic pitch, total pitch, flapping angle, number of blade airfoil segments and spanwise position, number of blade elements, source term sub-relaxation factor and other rotor geometric information and AASM control information.

[0086] B3. Read and configure mesh information. This program implements the reading of Fluent msh format (.msh / .cas) mesh files as its main supported mesh file format. In addition, it also supports the reading and conversion of CGNS (.cgns), CFL3D (.fmt / .xyz), GMsh (.msh) and other format files. Write mesh index information for memory allocation.

[0087] B4. AASM pre-processing, initialize the propeller disk area. First, directly query the internal grid block index of the propeller disk through the propeller disk surface index information, or search the grid cells inside the propeller disk through the propeller disk geometric parameters, and search and store the grid cells that will participate in the subsequent AASM calculation. According to the time step information, the blade area of ​​each time step position is individually marked. Momentum source term memory space is allocated to the grid cells participating in the AASM calculation.

[0088] B5. Flow field initialization: use the incoming flow parameters to directly assign values ​​to the flow field in the computational domain for initialization.

[0089] B6. CFD flow field solution, according to the solution configuration and control parameters of the corresponding grid block, the double time stepping method is used for iterative calculation. In the flow field calculation, the U-RANS momentum equation is used as the main control equation, the finite volume method is used to divide the calculation domain into a series of control volumes, and the partial differential equation is discretized by the integral form of the conservation equation. The three-dimensional unsteady compressible flow field control equation after adding the source term is as follows:

[0090]

[0091] Where V is the volume of the control body, S is the surface area of ​​the control body, n is the normal vector outside the control body, and the conserved variables W and inviscid flux F are c , viscous flux F v , the source term S is:

[0092] Absolute rotor speed q = [u, v, w] T , involving angular velocity q ω =[u ω ,v ω ,w ω ] T , ρ, p, e, h represent the density, pressure, internal energy and enthalpy of the fluid unit respectively, τ is the viscous stress, Θ is the work done by the viscous force and heat conduction on the unit, i x 、i y 、i z Represent the unit vectors in three directions respectively.

[0093] For the solution of unsteady flow field, the dual-time method is needed to advance the solution, which has second-order accuracy in time. The internal sub-iteration is based on the efficient implicit LU-SGS (Lower-Upper Symmetric Gauss-Seidel) algorithm, introducing pseudo-time τ, and the pseudo-time derivative term preprocessed with low Mach number will be introduced into the conservation quantity. The expression is as follows:

[0094]

[0095] Among them, Q is the pseudo time step conservation variable, and Γ is the low Mach number preconditioning matrix.

[0096] In order to improve the accuracy of flow field capture, the Roe inverse wind format is used for spatial discretization. According to the Roe format alone, the inviscid flux F(W) on the grid interface ij is ij The calculation formula is:

[0097]

[0098] Among them, L and R represent the left and right units respectively, |A Roe | ij is the Roe average Jacobian matrix.

[0099] The two-equation K-Epsilon (turbulent kinetic energy-turbulent dissipation rate) turbulence model is used to close the U-RANS momentum equation to improve the ability of the U-RANS momentum equation to capture unsteady flows. The turbulent eddy viscosity μ t The calculation is as follows:

[0100] μ t =ρC μ f μ kT (11).

[0101] Where ρ is the density, C μ is the model coefficient, f μ is the damping function, k is the turbulent kinetic energy, and T is the turbulent time scale.

[0102] The transmission equations of turbulent kinetic energy k and turbulent dissipation rate ε are:

[0103]

[0104] in, is the average velocity, μ is the dynamic viscosity, σ k , σ ε , C ε1 and C ε2 is the model coefficient, P k and P ε is the result term, f2 is the damping function, S k and S ε is the source term, and ε0 is the ambient turbulence value in the source term that offsets the turbulence attenuation.

[0105] In this embodiment, to obtain the U-RANS momentum equation, it is necessary to decompose each solution variable φ in the instantaneous NS equation into its average value and its fluctuation component φ′, expressed as:

[0106]

[0107] In this embodiment, the averaging can be considered as the time average of the steady-state case and the overall average of the repeatable transient case. The average mass, momentum and energy transfer equations are:

[0108]

[0109] Where ρ is the density, is the average speed, is the average pressure, I is the unit tensor, is the average viscous stress tensor, f b is the resultant of the body forces, is the average total energy per unit mass, is the average heat flux.

[0110] The additional term that appears in the momentum and energy transfer equations is the stress tensor T RANS , defined as follows:

[0111]

[0112] Where k is the turbulent kinetic energy.

[0113] In the time discretization process of each pseudo-iteration step, the velocity field and density field of the blade area are extracted and used as the input of the subsequent AASM calculation.

[0114] (3) Figure 3 The calculation demonstration diagram of the AASM method is shown. The AASM calculation and CFD calculation are coupled, the blade area velocity, density field input interface and momentum source term output interface are established, and the effect of the blade on the flow field is calculated by AASM and added to the CFD solution. The specific steps include:

[0115] C1. Establish coordinate transformation relationship. Figure 4 As shown in the figure, the expression of the transformation matrix M1 from the global coordinate system (x, y, z) to the propeller disc coordinate system (ξ, η, ζ) is:

[0116]

[0117] Among them, A and B represent the pitch angle and left pitch angle of the propeller disk respectively, (ξ,η,ζ) is the rectangular coordinate system of the propeller disk, (x,y,z) is the global rectangular coordinate system, (x c ,y c ,z c ) is the global coordinate of the rotor's center of rotation.

[0118] like Figure 5 As shown in Figure 2, the expression of the transformation matrix M2 from the propeller disc coordinate system (ξ, η, ζ) to the blade coordinate system (C, R, Z) is:

[0119]

[0120] Where ψ is the blade azimuth, β is the blade flapping angle, (e C ,e R ,e Z ) is the unit vector of the blade coordinate system, (e ξ ,e η ,e ζ) is the unit vector of the propeller disc rectangular coordinate system, sgn is the sign function, which is related to the rotation direction of the rotor and its expression is:

[0121]

[0122] C2. The velocity field extracted from the CFD solution is distributed to the blade element using the three-dimensional integral sampling method. The velocity field v and density field ρ around the blade element are convolved using the three-dimensional anisotropic Gaussian kernel function G and multiplied by the unit volume V. cell The weighted average G norm , eliminating the ambiguity of the point sampling position. The physical quantity of each blade element control point contains the information of the flow field and wake around the entire blade, and there is no need to correct the bound loop quantity. The form of the kernel function will be given when introducing the blade element force distribution. The blade element velocity field U e and the density field ρ e The calculation formula is as follows:

[0123]

[0124] Among them, (c e ,r e ,z e ) are the coordinates of the blade element center in the blade coordinate system (O, R, Z), and (c, r, δ) are the coordinates of the volume unit in the blade coordinate system (O, R, Z).

[0125] C3. Use blade element theory to calculate blade element force. Convert the velocity components (u, v, w) in the global coordinate system to the velocity components (u ξ ,v η ,w ζ ). Select the blade profile at any azimuth angle ψ and radius r, and consider the first-order flapping term of the blade β=β0-β 1c cosψ-β 1s sinψ, where β0 is the propeller disk taper angle, β 1c is the rear chamfer of the propeller disc, β 1s is the chamfer of the blade side. Take into account the blade circumferential rotation speed Ωr and blade flapping speed The velocity component (V C ,V R ,V Z ), expressed as:

[0126]

[0127] According to the velocity component V of the blade section Z , V C , blade twist angle θ0, total pitch θ 75 , lateral periodic pitch θ 1cand longitudinal periodic pitch θ 1s , defines the pitch angle Inflow angle β * and effective angle of attack α e :

[0128]

[0129] From the c81 table of airfoil aerodynamics, the lift coefficient C corresponding to the angle of attack and Mach number of the section is obtained by two-dimensional Lagrangian interpolation. l and the drag coefficient C d , then the lift dL and drag dD of the blade element micro-segment are:

[0130]

[0131] in, c is the blade element chord length, and dr is the blade element micro-segment length.

[0132] Then, if Figure 6 As shown, the lift and drag of the blade section are used to calculate the pull dT and drag dQ perpendicular to the propeller disc plane, expressed as:

[0133]

[0134] Since there is no aerodynamic force in the blade span direction, the blade aerodynamic force F in the global coordinate system is obtained by coordinate transformation. x ,F y ,F z ), expressed as:

[0135]

[0136] C4. Convolve the blade element force with the Gaussian kernel function and convert it into a momentum source term distributed to the volume grid unit. An anisotropic three-dimensional Gaussian kernel is defined with the following form:

[0137]

[0138] The source terms are distributed in the chord direction and thickness direction of the blade section according to the 3σ criterion, which makes the influence of the blade element force at the leading and trailing edges of the blade decay to 0. The smearing factors in the chord direction and thickness direction are defined respectively. and In order to improve the spanwise continuity of the blade element force distribution, the influence range of the local blade element force will be expanded to the three surrounding blade elements. At the same time, the blade element itself will also be affected by the nearby blade elements. The spanwise smearing parameter ε is defined r =κ(R tip -R root ) / n, where R tip , R rootare the tip and root positions of the blade, n is the number of blade segments, and an appropriate spanwise load correction coefficient κ=2 is selected. The actuating surface model after the introduction of the correction coefficient κ will not affect the original spanwise load of the blade, and can simulate the three-dimensional effect of the blade tip, thereby replacing the blade root and blade tip loss model. The introduction of these three smearing parameters increases the convergence of the calculation and solves the problem of poor flow field stability of the traditional momentum source method.

[0139] The aerodynamic force of each blade element calculated by C3 is then convolved with a three-dimensional anisotropic Gaussian kernel and the grid cell volume V is taken into account in the normalization process. cell The influence of , to balance the fluctuation of the grid size, finally obtains the force vector f on the volume element ASM , the calculation formula is as follows:

[0140]

[0141] In the current iteration step, the force exerted by the blade on the fluid micro-group is the unsteady momentum source term S:

[0142] S=-f ASM (27).

[0143] In order to further improve the convergence of the flow field, this embodiment introduces the sub-relaxation factor α of the unsteady momentum source term. According to the source term S of the previous iteration step, i-1 The source term S calculated for the current iteration step i For relaxation, the source term sub-relaxation factor α is generally taken as 0.5, then:

[0144] S i =α(S i -S i-1 )+S i-1 (28).

[0145] Finally, the source term is added to the residual calculation of the U-RANS momentum equation to complete the coupling closed loop of the AASM method and CFD simulation.

[0146] C5. Continue to execute the CFD solution process. When the calculation converges or reaches the maximum number of pseudo-iteration steps, enter the next time step calculation and update the blade area index of the next time step to map the actual blade position after the rotor rotates.

[0147] C6. Repeat the coupling process of CFD solution and AASM solution, and output the aerodynamic characteristics of the rotor, single blade, and single blade section in real time during the calculation process, including aerodynamic load, induced speed, aerodynamic coefficient, blade-vortex interference position, etc. If there is a fuselage, the pressure data of the fuselage surface and the specified surface monitoring point can also be output. Until the maximum physical time or maximum time step is reached, all the conserved variable data of the flow field in the calculation domain are output. After obtaining the rotor aerodynamic characteristics and flow field data, the rotor aerodynamic analysis can be carried out, and it can be used as input to further carry out complex full-machine interference and maneuvering flight state calculation research work with extremely high resource consumption.

[0148] This embodiment couples the AASM method with the CFD simulation method and applies it to high-fidelity flow field calculation of a rotorcraft. Figure 7 A schematic diagram showing the variation of the tension coefficient with the azimuth angle during the tests of the AASM method, the traditional unsteady momentum source method, and the advanced actuation surface method provided in this embodiment; Figure 8 Schematic diagram of the lift variation with azimuth angle when r / R=0.50 for AASM method, traditional unsteady momentum source method and moving nested grid method; Fig. 9 Schematic diagram of the lift variation with azimuth angle when r / R=0.65 for AASM method, traditional unsteady momentum source method and moving nested grid method; Fig.10 Schematic diagram of the lift variation with azimuth angle when r / R=0.89 for AASM method, traditional unsteady momentum source method and moving nested grid method; Fig.11Schematic diagram of the lift variation with azimuth angle when r / R=0.99 for the AASM method, the traditional unsteady momentum source method and the moving nested grid method. Compared with the traditional moving nested grid method, there is no need to model and generate grids for the real blade model, which reduces the number of grids in the computational domain and avoids the time overhead of nested grid digging, donor search and interpolation in CFD calculations. Compared with the traditional unsteady momentum source method, there is no need to search the blade area in each CFD solution iteration step, which avoids the time overhead caused by regional search; the AASM method is used to calculate and distribute the momentum source term, without the need for additional root / tip loss models and virtual interpolation grids, which reduces the computational cost and improves the computational convergence of the blade edge area and the continuity of the rotor flow field; three-dimensional integral sampling is used when extracting the blade velocity field and density field, which eliminates the influence of the sampling position on the calculation, does not require bound circulation correction, and improves the calculation accuracy of the blade aerodynamic force. Moreover, compared with flow field simulation software, the computational efficiency of this embodiment in rotorcraft flow field simulation is 40 times that of its motion nested grid method and 6 times that of its time-accurate virtual disk method based on the momentum source model. Under the same set of grids, the rotor flow field calculated by this embodiment has better continuity and convergence, and the precision of the obtained rotor blade aerodynamic load is greatly improved. The efficient computational method makes the computation of complex rotor flow field, which originally consumes a lot of computational resources, feasible.

[0149] In view of the task requirements of rotorcraft aerodynamic characteristics analysis, this embodiment breaks through the key technology of high-fidelity rotor flow field rapid evaluation, proposes an efficient flow field numerical simulation technology for rotorcraft aerodynamic characteristics analysis, couples the AASM method with the CFD simulation method based on the U-RANS momentum equation, creates a general hybrid grid generation strategy suitable for the rotor momentum source method, and applies it to the unsteady flow field calculation of the rotorcraft to obtain accurate rotor unsteady aerodynamic loads, with higher computational efficiency, flow field continuity and convergence, and is suitable for rotorcraft flow field simulation under most flight conditions, which not only fills the current shortcomings in the numerical simulation algorithm of the unsteady momentum source of the rotorcraft, but also solves the problem that the blade aerodynamic load calculated by the traditional momentum source method fluctuates greatly, and it is impossible to further carry out the research on rotor aerodynamic noise, complex whole-machine interference and maneuvering flight state. At the same time, this embodiment also reduces the difficulty of rotor flow field simulation, and solves the problem that the traditional rotor flow field numerical simulation algorithm requires the generation of high-resolution grids, resulting in very high computational costs.

[0150] In an exemplary embodiment, a computer device is also provided, including a memory and a processor, wherein a computer program is stored in the memory, and the processor implements the above-mentioned numerical simulation method of the rotorcraft flow field when executing the computer program.

[0151] In an exemplary embodiment, a computer-readable storage medium is provided, storing a computer program, which implements the above-mentioned rotorcraft flow field numerical simulation method when executed by a processor.

[0152] In an exemplary embodiment, a computer program product is provided, including a computer program, which implements the above-mentioned rotorcraft flow field numerical simulation method when executed by a processor.

[0153] The technical features of the above embodiments may be combined arbitrarily. To make the description concise, not all possible combinations of the technical features in the above embodiments are described. However, as long as there is no contradiction in the combination of these technical features, they should be considered to be within the scope of this specification.

[0154] This article uses specific examples to illustrate the principles and implementation methods of this application. The description of the above embodiments is only used to help understand the method and core ideas of this application. At the same time, for those skilled in the art, according to the ideas of this application, there will be changes in the specific implementation methods and application scope. In summary, the content of this specification should not be understood as limiting this application.

Claims

1. A method for numerical simulation of the flow field of a rotorcraft, characterized in that: The rotorcraft flow field numerical simulation method comprises: Based on the rotorcraft under given flight conditions, a rotorcraft model is established; the rotorcraft model is a three-dimensional CAD model including an aircraft body model and a rotor blade model; Based on the rotorcraft model, a hybrid grid system is generated; the hybrid grid system includes a grid structure corresponding to the aircraft body model and a grid structure corresponding to the rotor blade disc model, and the hybrid grid system is used to provide a grid structure and boundary conditions required for flow field calculation; According to the hybrid grid system, a flow field CFD numerical simulation is performed on the rotorcraft to obtain conservation variable information of the blade region; the conservation variable information includes velocity field information and density field information of the blade region; According to the conserved variable information of the blade area, the unsteady momentum source term of the blade area is calculated by using the AASM method; According to the unsteady momentum source term of the blade area, the aerodynamic characteristic data and flow field data of the rotor under a given flight condition are determined.

2. The method for numerical simulation of rotorcraft flow field according to claim 1, characterized in that: Based on the rotorcraft under given flight conditions, a rotorcraft model is established, including: For the rotorcraft under given flight conditions, according to the basic dimensions of the rotorcraft, CAD software is used to respectively establish the aircraft body model and the rotor disc model corresponding to the rotorcraft to obtain the rotorcraft model.

3. The method for numerical simulation of rotorcraft flow field according to claim 1, characterized in that: Based on the rotorcraft model, a hybrid grid system is generated, specifically including: Based on the aircraft body model and the background grid area, an unstructured tetrahedral grid block is generated by adopting a front advancement method; the unstructured tetrahedral grid block is a grid structure corresponding to the aircraft body model, and the unstructured tetrahedral grid block includes a plurality of tetrahedral grid units; Based on the rotor disc model and the rotor disc region, a hexahedral grid block is generated by using a sweeping method; the hexahedral grid block is a grid structure corresponding to the rotor disc model, and the hexahedral grid block includes a plurality of hexahedral grid units; The surface mesh units corresponding to the external flow field and the surface mesh units corresponding to the disk flow field are topologically shared at the disk interface, and pyramid mesh units are used to smoothly transition the tetrahedral mesh units and the hexahedral mesh units; wherein the tetrahedral mesh units, the hexahedral mesh units, the surface mesh units and the pyramid mesh units are different mesh types in the hybrid mesh system.

4. The method for numerical simulation of rotorcraft flow field according to claim 1, characterized in that: According to the hybrid grid system, the flow field CFD numerical simulation of the rotorcraft is performed to obtain the conservation variable information of the blade area, which specifically includes: Based on the given flight conditions and incoming flow conditions, the flow field CFD numerical simulation of the rotorcraft is performed according to the hybrid grid system, the U-RANS momentum equation is used as the main control equation, and the U-RANS momentum equation is solved by time advancement to obtain the velocity field information and density field information of the blade area in each iteration step.

5. The method for numerical simulation of rotorcraft flow field according to claim 4, characterized in that: According to the conserved variable information of the blade area, the unsteady momentum source term of the blade area is calculated by using the AASM method, which specifically includes: Performing three-dimensional integral sampling on the velocity field information and the density field information of the blade area in each iteration step to obtain a three-dimensional integral sampling result; The three-dimensional integral sampling result is used as the input of the AASM method to construct a velocity triangle on each blade section of the blade area; the velocity triangle is used to determine the blade element force of the blade section; According to the velocity triangle on each blade section, based on the blade element theory, the blade element force of each blade section is calculated to obtain a blade element force calculation result; The blade element force calculation results of each blade section are converted into momentum source terms, and are convolved with anisotropic three-dimensional Gaussian kernels to obtain unsteady momentum source terms in the blade area.

6. The method for numerical simulation of rotorcraft flow field according to claim 5, characterized in that: According to the unsteady momentum source term of the blade area, the rotor aerodynamic characteristic data and flow field data under a given flight condition are determined, specifically including: The unsteady momentum source term of the blade region is added to the residual calculation of the U-RANS momentum equation for iterative calculation, and the rotor aerodynamic characteristic data and the flow field data are output in real time.

7. The method for numerical simulation of rotorcraft flow field according to claim 1, characterized in that: The rotor aerodynamic characteristic data include aerodynamic load, induced speed, aerodynamic coefficient and blade-vortex interference position; the flow field data include all conservation variable data in the flow field calculation process.

8. A computer device comprising: A memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor executes the computer program to implement the method for numerical simulation of the rotorcraft flow field according to any one of claims 1 to 7.

9. A computer-readable storage medium having a computer program stored thereon, characterized in that: When the computer program is executed by a processor, the method for numerical simulation of the rotorcraft flow field according to any one of claims 1 to 7 is implemented.

10. A computer program product, comprising a computer program, characterized in that When the computer program is executed by a processor, the method for numerical simulation of the rotorcraft flow field according to any one of claims 1 to 7 is implemented.

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