Method for calculating unsteady aerodynamic interference of multi-component combined aircraft

Abstract

The invention provides a multi-component combined aircraft unsteady aerodynamic interference calculation method, which comprises the following steps that: a plurality of combinations are determined, and each combination comprises at least one component; carrying out grid division on each component; dividing inner and outer layers of background grids for all combinations; setting calculation parameters for each combination, wherein the calculation parameters comprise the forward flight speed, the density, the number of calculation turns, the physical time step of each turn and the number of pseudo time steps of each turn; balancing calculation parameters and balancing targets are set for the combination comprising the rotor components; and calculating an interference flow field of each combination, and obtaining a rule and a mechanism of unsteady interference suffered by the component. The method comprehensively considers the influence of the multi-component combination of the aircraft, can accurately simulate and analyze the aerodynamic interference characteristics of each component of the aircraft in a real working environment, and provides support for aerodynamic layout design and aerodynamic interference characteristic analysis of the aircraft. The unsteady aerodynamic force obtained through calculation can provide load input for dynamics calculation, and multidisciplinary coupling analysis is achieved.

Description

TECHNICAL FIELD

[0001] The present application belongs to the field of aircraft technology, and particularly relates to a method for calculating unsteady aerodynamic interference of a multi-component combined aircraft. BACKGROUND

[0002] The current aircraft of helicopters or new configuration high-speed helicopters has multiple rotating components (rotors, tail rotors, etc.) and static components (fuselages, horizontal tails, vertical tails, wings, etc.). In the flight process, the flow field presents obvious unsteady characteristics due to the existence of rotating components. Taking a conventional helicopter as an example, in the hovering and low-speed forward flight state, the fuselage and tail rotor are almost completely in the wake interference flow field of the rotor. This interference will generate impact load and periodic load on the helicopter fuselage and tail rotor, thereby affecting the handling and stability characteristics, vibration and noise characteristics of the helicopter. Similarly, in the high-speed forward flight state, although the vortex wake does not directly act on the fuselage and tail rotor, the induced flow field generated by the rotor still produces obvious interference on the fuselage and tail rotor. In addition, the rotor will also be disturbed by the flow field changes generated by the fuselage and tail rotor, resulting in changes in the rotor disc load. The mutual interference between components leads to a large difference between the performance of each component and the isolated component, which greatly affects the accurate evaluation of the helicopter performance. Therefore, the calculation of the aerodynamic interference characteristics of the helicopter under the interference state of the full machine (multiple rotors, fuselages, tail rotors or propulsion rotors) can provide accurate performance data for the evaluation of the flight performance of the helicopter, and has important practical significance for analyzing the performance of the full machine and the interference mechanism of the components.

[0003] The existing calculation and analysis of aerodynamic interference characteristics of aircraft are mostly limited to the momentum source method, which cannot accurately simulate the real working environment of the aircraft, and the analysis of some complex interference problems is not thorough enough. In addition, the traditional CFD calculation has the problem of low calculation efficiency and cannot be automatically trimmed. SUMMARY

[0004] To solve the problem of low calculation efficiency and the problem of not being able to automatically trim in the traditional CFD calculation in the prior art, the present application proposes a CFD method based on two-layer background grids and automatic trimming, which greatly improves the calculation efficiency and realizes the calculation of unsteady aerodynamic interference of a multi-component combined aircraft. Based on the moving nested grid, a method for calculating unsteady aerodynamic interference of a multi-component combined aircraft in a trimmed state is provided, which provides support for aerodynamic layout design, aerodynamic interference characteristic analysis and full machine dynamics analysis of the aircraft. The technical solution is as follows: In a first aspect, a method for calculating unsteady aerodynamic interference of a multi-component combined aircraft is provided, and the method comprises the following steps: S1, determining a plurality of combinations, each combination comprising at least one component: S2, performing mesh division of each component; S3, divide the inner and outer two layers of background grid for all combinations; S4, set the calculation parameters for each combination, the calculation parameters including the forward speed, density, calculation number of circles, physical time step of each circle, and pseudo time step number of each circle; S5, set the trimming calculation parameters and trimming targets for the combination containing the rotor component; S6, calculate the interference flow field of each combination, and obtain the unsteady interference law and mechanism suffered by the component, the interference flow field including the aerodynamic force, velocity field, and vorticity field.

[0005] Optionally, 2. The method according to claim 1, wherein in S1, the combination mode with the most aircraft components is determined, and then other combinations are determined according to the characteristics of the aircraft components, each combination mode including at least one component; the components are divided into moving components and static components.

[0006] Optionally, 3. The method according to claim 1, wherein in S2, the static components are subjected to local clipping processing and the moving components are subjected to root cutting processing for each component in S1; and then each component is subjected to grid division, and the grid type of each component is a structure grid or a non-structure grid. For the moving component, the outer boundary is set to be 1.5 times to 1.8 times the profile chord length, and the first layer thickness of the blade boundary layer grid is set to be in the order of magnitude of the corresponding chord length*10^-5; the surface grid of the moving component blade leading edge and blade tip is encrypted. For the static component, the outer boundary is set to be 1.2 times to 1.5 times the reference static component length, and the first layer thickness of the static component boundary layer grid is set to be in the order of magnitude of the reference static component length*10^-5; the surface grid of the static component leading edge is encrypted.

[0007] Optionally, 4. The method according to claim 1, wherein in S3, the inner background grid is divided to contain all the components, the outer boundary of the inner background grid is 1.5 times to 2.0 times the maximum length of each component in each direction, and the inner background grid is encrypted in the complex flow area of each component; the outer boundary of the outer background grid is set to be 5 times to 10 times the length of the longest component in each direction; and the grid size ratio at the data exchange place of the outer background grid and the inner background grid is 1 times to 1.2 times.

[0008] Optionally, 5. The method according to claim 1, wherein in S4, For the combination containing the rotor component, the calculation parameters further include the number of rotors, the total number of rotor blades, the rotor tip speed, the rotation direction, and the rotor rotation center. For combinations that include propeller components, the calculation parameters also include the number of propellers, the total number of propeller blades, the propeller tip speed, the direction of propeller rotation, the propeller's axis of rotation, and the center of propeller rotation. For a combination that includes at least one of a fuselage and a wing component, the calculation parameters also include fuselage / wing attitude and fuselage / wing attitude reference position; Set up multi-level motion relationships for combinations containing moving parts; set up multi-level motion relationships for moving parts: first perform blade flapping motion and pitch change motion, then perform rotational motion; When realizing the flapping and pitch-changing motion of the propeller blades, a rotating coordinate system is established with the rotation center of the moving part as the origin. The X direction is the incoming flow direction, the Y direction is vertically downward, and the Z direction is determined by the right-hand rule. In the rotating coordinate system, the motion of the moving part considers the rotation, flapping, and pitch-changing motion of the rotor / propeller. The flapping and pitch-changing laws are expressed as functions of the azimuth angle.

[0009] in, It is the blade azimuth angle. The coefficient of motion during the swing. The pitch motion coefficient.

[0010] When achieving rotational motion, the rotational law is a function related to the tip velocity of the moving part.

[0011] Optionally, 6. The method according to claim 1, characterized in that, in S5, Set the trim calculation parameters, including: initial number of cycles, azimuth angle calculated in a single iteration, total number of trim iterations, trim target quantity, damping factor, and control difference calculated by Jacobi matrix; The trim targets set include four quantities: drag, lift, pitching moment, and rolling moment. Alternatively, the trim target can include three quantities: lift, pitch moment, and roll moment.

[0012] Optionally, 7. The method according to claim 1, wherein S6 specifically comprises: S61. For each combination, read in the inner and outer background meshes and the calculation parameters; For combinations that include rotor components, it is also necessary to read in the rotor mesh, trim calculation parameters, and trim objectives; For assemblies that include propeller components, the propeller mesh also needs to be read in; For combinations that include fuselage / wing components, the fuselage / wing mesh also needs to be read in; S62. Perform hole-punching and contribution cell search on the inner background mesh and the outer background mesh, and store the nesting relationship between the two in an array; S63. Initialize the flow field information based on the forward velocity and density in the calculation parameters: assign the forward velocity and density variables to the grid divided in S2; For combinations that include rotor components, the forward velocity and density variables are assigned to the rotor mesh; For combinations that include propeller components, the forward velocity and density variables are assigned to the propeller mesh; For combinations that include fuselage / wing components, the forward velocity and density variables are assigned to the fuselage / wing mesh; S64. Calculate the disturbance flow field based on the number of calculation revolutions, physical time steps per revolution, and pseudo time steps per revolution set in S4: S65. Based on the disturbance flow field, obtain the aerodynamic force, velocity field, and vorticity field of the component under different physical time steps for each combination; compare and analyze the aerodynamic force, velocity field, and vorticity field of the component under different combinations to obtain the law and mechanism of unsteady disturbance to the component.

[0013] Alternatively, S64 specifically includes: For combinations containing rotor components, the final control matrix is ​​calculated based on the trim calculation parameters and trim objectives. The three-dimensional unsteady Reynolds-averaged NS (RANS) equations are solved based on the number of calculation revolutions, the physical time step per revolution, and the pseudo time step per revolution. Information of each component is exchanged through nesting relationships, and information is exchanged between the inner background mesh and the outer background mesh to finally obtain the disturbance flow field. The formula for balancing is:

[0014] in, The target quantity matrix of the rotor. This represents the rotor control matrix. The damping factor is 0 < <1, The Jacobian matrix calculated for balancing, This represents the difference in the target quantity matrix. This indicates the manipulation matrix difference, and the Jacobian matrix is ​​obtained through iterative calculation of the difference manipulation. For a combination that includes a propeller component, or a combination that includes at least one of a fuselage and a wing component, the three-dimensional unsteady Reynolds-averaged NS (RANS) equations are solved based on the number of revolutions, the physical time step per revolution, and the pseudo time step per revolution. Information about each component is exchanged through nesting relationships, and information is exchanged between the inner background mesh and the outer background mesh, ultimately yielding the disturbance flow field.

[0015] The beneficial effects of this invention are at least as follows: A comprehensive aerodynamic characteristic analysis method considering the combined interference of multiple aircraft components is proposed. Based on CFD numerical simulation with a nested motion mesh and combined with a two-layer background mesh technique, the method significantly improves computational efficiency. By implementing combined interference schemes, processing computational models, dividing computational meshes, defining the motion of each component, and setting up numerical solutions, the method comprehensively considers the influence of the combination of multiple aircraft components and accurately simulates and analyzes the aerodynamic interference characteristics of each component under the real operating environment of the aircraft. This provides support for the aerodynamic layout design and aerodynamic interference characteristic analysis of the aircraft. Simultaneously, the calculated unsteady aerodynamic forces can provide load inputs for dynamic calculations, enabling multidisciplinary coupled analysis. Attached Figure Description

[0016] Figure 1 A flowchart for calculating aerodynamic interference of a multi-component combined aircraft provided in an embodiment of the present invention; Figure 2 This is a schematic diagram of the overall shape of the machine provided in an embodiment of the present invention; Figure 3 This is a schematic diagram of the assembly of various components provided in an embodiment of the present invention; Figure 4 A schematic diagram of a two-layer background grid provided in an embodiment of the present invention; Figure 5 A schematic diagram of the Q isosurface of a multi-component aircraft provided for an embodiment of the present invention (Q=0.005 isosurface during forward flight at 400km / h). Figure 6 A schematic diagram of the induced velocity field for different component combinations provided in the embodiments of the present invention (induced velocity in the Y direction at position Y=0.1m (400km / h)). Figure 7 This is a schematic diagram of the unsteady aerodynamics of different components provided in the embodiments of the present invention; Figure 8 A schematic diagram showing the aerodynamic changes of the left propeller under different combination states, provided in an embodiment of the present invention; Figure 9 The spectrum diagram of the left propeller thrust under full-engine conditions is provided for an embodiment of the present invention. Detailed Implementation

[0017] 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.

[0018] 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.

[0019] 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.

[0020] The present invention will now be described in detail with reference to the accompanying drawings and specific embodiments.

[0021] The present invention will now be described in further detail with reference to specific embodiments and accompanying drawings.

[0022] This invention provides a method for calculating unsteady aerodynamic disturbances in multi-component aircraft. Please refer to [link to relevant documentation]. Figure 1 This includes the following steps: S1, determine several combinations, each combination including at least one component; In S1, the combination with the most aircraft components is determined (i.e., the whole aircraft combination). Then, other combinations are determined based on the characteristics of the aircraft components (rotors / fuselage / propellers, etc.). Each combination includes at least one component, and all combinations are determined. The components are mainly divided into moving components and stationary components. This invention can realize the combination of multiple moving components and multiple stationary components. For example, the moving components are no more than 20 rotors, no more than 20 propellers, and no more than 10 stationary components (fuselage / wings, etc.).

[0023] S2, perform mesh generation for each component; For each component in S1, based on the principle of not affecting the basic flow characteristics of the model, local trimming is performed on static components, and root cutting is performed on the non-airfoil segments at the root of dynamic components such as rotors and propellers to avoid interference problems in the component meshes. Each component in S1 is meshed, and the mesh type for each component is either a structured mesh or an unstructured mesh. For moving parts (rotor / propeller), the outer boundary is set to 1.5-1.8 times the section chord length, and the thickness of the first layer of the blade boundary layer mesh is set to the order of the corresponding chord length * 10^-5; the surface mesh of the blade leading edge and tip of the moving part is refined. For stationary components (fuselage / wing), set the outer boundary to 1.2-1.5 times the length of the reference stationary component, and set the thickness of the first layer of the boundary layer mesh of the stationary component to be on the order of the length of the reference stationary component * 10^-5; refine the surface mesh of the leading edge mesh of the stationary component.

[0024] S3, divides all combinations into inner and outer background grids; The inner background mesh is divided to include all components. The outer boundary of the inner background mesh is generally 1.5 to 2.0 times the maximum length of each component in each direction, such as 1.5 to 2.0 times in the X direction, 1.5 to 2.0 times in the Y direction, and 1.5 to 1.8 times in the Z direction.

[0025] In complex flow areas of each component, the inner background mesh is refined; The boundaries of the outer background mesh in each direction are generally set to be 5 to 10 times the length of the longest component; The ratio of the grid size at the data exchange point between the outer and inner background grids is generally 1 to 1.2 times to ensure that the nesting relationship between the background grids is stable and accurate.

[0026] S4, set the calculation parameters for each combination. The calculation parameters include forward flight speed, density, number of calculation laps, physical time steps per lap, and pseudo time steps per lap. The physical time steps per lap and pseudo time steps per lap can be set arbitrarily. S41. For combinations that include rotor components, the calculation parameters also include the number of rotors, the total number of rotor blades, the rotor tip speed, the direction of rotation, and the rotor rotation center. For combinations that include propeller components, the calculation parameters also include the number of propellers, the total number of propeller blades, the propeller tip speed, the direction of propeller rotation, the propeller's axis of rotation, and the center of propeller rotation. For a combination that includes at least one of fuselage and wing components, the calculation parameters also include fuselage / wing attitude (pitch and roll angles) and fuselage / wing attitude reference position; S42. For combinations containing moving parts, it is further necessary to set up multi-level motion relationships; Set up a multi-stage motion relationship for the moving parts: first perform blade flapping motion and pitch change motion, then perform rotational motion; When realizing the flapping and pitch-changing motion of the propeller blades, a rotating coordinate system is established with the rotation center of the moving component as the origin. The X direction is the incoming flow direction, the Y direction is vertically downward, and the Z direction is determined using the right-hand rule. In this rotating coordinate system, the motion of the moving component can be considered as the rotation, flapping, and pitch-changing motion of the rotor / propeller. The flapping and pitch-changing laws can be expressed as functions of the azimuth angle, such as:

[0027] in, It is the blade azimuth angle. The coefficient of motion during the swing. The pitch motion coefficient.

[0028] When achieving rotational motion, the rotational law is a function related to the tip velocity of the moving part in S41.

[0029] S5, for combinations that include rotor components, further set the trim calculation parameters and trim targets; Set the balancing calculation parameters: initial number of cycles, azimuth angle calculated in a single iteration, total number of balancing iterations, balancing target quantity, damping factor, and manipulation difference calculated using the Jacobian matrix.

[0030] The trim targets set include four quantities: drag, lift, pitching moment, and rolling moment. Alternatively, the trim target can include three quantities: lift, pitch moment, and roll moment.

[0031] S6 calculates the disturbance flow field for each combination and obtains the unsteady disturbance law and mechanism of the component. The disturbance flow field includes aerodynamic force, velocity field and vorticity field.

[0032] S61. For each combination, read in the inner and outer background meshes of S3, and read in the calculation parameters of S4. For combinations that include rotor components, it is also necessary to read in the rotor mesh of S2 and the trim calculation parameters and trim target of S5; For assemblies that include propeller components, the propeller grid from S2 also needs to be read in; For combinations that include fuselage / wing components, the fuselage / wing grid from S2 also needs to be read in; S62. Perform hole-punching and contribution cell search on the inner background mesh and the outer background mesh, and store the nesting relationship between the two in an array; For combinations that include fuselage / wing components, it is necessary to further perform hole punching and contributing cell search on the fuselage / wing component mesh and the inner background mesh, and store the nesting relationship between the two in an array; S63. Based on the forward velocity and density in the calculation parameters set in S4, initialize the flow field information: assign the forward velocity and density variables to the grid divided in S2; For combinations that include rotor components, the forward velocity and density variables are assigned to the rotor mesh; For combinations that include propeller components, the forward velocity and density variables are assigned to the propeller mesh; For combinations that include fuselage / wing components, the forward velocity and density variables are assigned to the fuselage / wing mesh; S64. Calculate the disturbance flow field based on the number of calculation revolutions, physical time steps per revolution, and pseudo time steps per revolution set in S4: 1) For combinations containing rotor components, the final control matrix is ​​calculated based on the trim calculation parameters and trim target set in S5. The three-dimensional unsteady Reynolds-averaged NS (RANS) equations are solved based on the number of calculation revolutions, the physical time step per revolution, and the pseudo time step per revolution. The information of each component is exchanged through nesting relationships, and the inner background grid is exchanged with the outer background grid to finally obtain the disturbance flow field.

[0033] The formula for balancing is as follows:

[0034] in, The target quantity matrix of the rotor. This represents the rotor control matrix. (0< <1) is the damping factor. The Jacobian matrix calculated for balancing, This represents the difference in the target quantity matrix. This represents the manipulation matrix difference. The Jacobian matrix is ​​obtained through iterative calculation using difference manipulation; 2) For combinations containing propeller components, the three-dimensional unsteady Reynolds-averaged NS (RANS) equations are solved based on the number of revolutions, physical time steps per revolution, and pseudo time steps per revolution. Information of each component is exchanged through nesting relationships, and information is exchanged between the inner background mesh and the outer background mesh to finally obtain the disturbance flow field. 3) For a combination that includes at least one of the fuselage and wing components, the three-dimensional unsteady Reynolds-averaged NS (RANS) equations are solved based on the number of calculation revolutions, the physical time step per revolution, and the pseudo time step per revolution. The information of each component is exchanged through nesting relationships, and the inner background mesh is exchanged with the outer background mesh to finally obtain the disturbance flow field. S65. Based on the disturbance flow field, obtain the aerodynamic force (six force elements), velocity field and vorticity field of each component under different physical time steps of each combination. By comparing and analyzing the aerodynamic forces (six force elements), velocity field, and vorticity field of the components under different combinations, the unsteady disturbance laws and mechanisms of the components are obtained.

[0035] See Figure 1 An embodiment of the present invention also provides a method for calculating unsteady aerodynamic disturbances of a multi-component aircraft, specifically including: S1, determine several combinations, each combination including at least one component: In this embodiment, the entire aircraft assembly consists of one rotor, one fuselage, and two propellers. Other components are combined in pairs or isolated components form a combination, resulting in the following combinations: isolated propeller, isolated rotor, isolated fuselage, rotor / fuselage, rotor / propeller, fuselage / propeller, and rotor / fuselage / propeller, a total of 7 combinations; see also Figure 2 and Figure 3 ; S2, perform mesh generation for each component: S21. Simplify the non-airfoil sections at the root of the rotor and tail rotor to avoid interference problems in the component mesh. Cut the rotor root to 0.2R and the propeller root to 0.3R. S22. Mesh the rotor / fuselage / propeller components, selecting the rotor and propeller blade mesh type as structured mesh and the fuselage mesh type as unstructured mesh; The rotor radius R is 10m, the propeller radius R is 1.905m, the outer boundary of the blade grid is set to 1.5 times the section chord length, and the thickness of the first layer of the blade boundary layer grid is set to the order of the corresponding chord length * 10^-5; the grid of the rotor / propeller blade leading edge and tip is refined; The reference length of the fuselage is 8.18m, the outer boundary is set to 12.27m, the thickness of the first layer of the fuselage boundary layer mesh is 8.18*10^-5m, and the surface mesh of the fuselage leading edge is refined. S3, divides all combinations into inner and outer background meshes: The inner background mesh is divided to include all components, with outer boundaries of -15m and 20m in the X direction, -15m and 20m in the Y direction, and -15m and 15m in the Z direction; Inner background mesh distribution: Spatial mesh refinement is performed on the inner background mesh in complex flow areas of the rotor / fuselage / propeller; The boundaries of the outer background mesh in each direction are set to -60m and 60m; The grid size ratio at the data exchange point between the outer and inner background grids is set to 1.1 to ensure a stable and accurate nesting relationship between the background grids. See also... Figure 4 .

[0036] S4. Set calculation parameters for each combination. The calculation parameters include forward flight speed of 400km / h, density of 1.225kg / m^3, number of calculation orbits of 5, physical time step of 0.5° per orbit, and pseudo time step of 10 per orbit. S41. For a combination that includes a rotor component, the calculation parameters also include 1 rotor, a total of 4 rotor blades, a rotor tip speed of 188m / s, a rotation direction of counterclockwise when viewed from above, and a rotor rotation center of (0, -0.15, 0). For the combination including the propeller component, the calculation parameters also include 2 propellers, a total of 10 propeller blades, a tip velocity of 216 m / s for both propellers, the left propeller rotates counterclockwise, the right propeller rotates clockwise (viewed from back to front), both propellers rotate around the X-axis, and the rotation center of the left propeller is (2.295, 2.85, -5.6), and the rotation center of the right propeller is (2.295, 2.85, 5.6). For combinations that include fuselage components, the calculation parameters also include fuselage pitch angle of 0°, roll angle of 0°, and fuselage reference position (0, 0, 0). S42. Set up multi-stage entanglement motions for the rotor and propeller. When realizing blade flapping and pitch-changing motions, establish a rotating coordinate system with the rotation center of the moving component as the origin. The X direction is the incoming flow direction, the Y direction is vertically downward, and the Z direction is determined using the right-hand rule. In this rotating coordinate system, the motion of the moving component can consider the rotation, flapping, and pitch-changing motions of the rotor / propeller. The flapping and pitch-changing laws can be expressed as functions of azimuth angle, such as:

[0037] It is the blade azimuth angle. The coefficient of motion during the swing. The variable pitch motion coefficient is the rotor collective pitch. The longitudinal periodic pitch is 7.3°. The lateral periodic pitch is 5.3°. The angle is -9°, the total pitch of the left and right propellers is 36°, and the other waving motion coefficients and pitch change motion coefficients are all 0.

[0038] S5, for combinations including rotor components, further set the trim calculation parameters and trim targets: Set the balancing calculation parameters as follows: initial number of laps is 3, azimuth angle for single iteration is 120°, total number of balancing iterations is 8, balancing target value is 3, damping factor is 0.6, and control difference calculated by Jacobi matrix is ​​0.2°. The trim target includes three quantities: lift, pitch moment, and roll moment, with corresponding coefficients of 0.007365, 0, and -0.00147, respectively.

[0039] S6, calculate the disturbance flow field for each combination, and obtain the unsteady disturbance law and mechanism of the component. The disturbance flow field includes aerodynamic force, velocity field, and vorticity field: S61. For each combination, read in the inner and outer background meshes of S3, and read in the calculation parameters of S4. For combinations that include rotor components, it is also necessary to read in the rotor mesh of S2 and the trim calculation parameters and trim target of S5; For assemblies that include propeller components, the propeller grid from S2 also needs to be read in; For combinations that include fuselage / wing components, the fuselage / wing grid from S2 also needs to be read in; S62. Perform hole-punching and contribution cell search on the inner background mesh and the outer background mesh, and store the nesting relationship between the two in an array; For combinations that include fuselage / wing components, it is necessary to further perform hole punching and contributing cell search on the fuselage / wing component mesh and the inner background mesh, and store the nesting relationship between the two in an array; S63. Based on the forward velocity and density in the calculation parameters set in S4, initialize the flow field information: assign the forward velocity and density variables to the grid divided in S2; For combinations that include rotor components, the forward velocity and density variables are assigned to the rotor mesh; For combinations that include propeller components, the forward velocity and density variables are assigned to the propeller mesh; For combinations that include fuselage / wing components, the forward velocity and density variables are assigned to the fuselage / wing mesh; S64. Calculate the disturbance flow field based on the number of calculation revolutions, physical time steps per revolution, and pseudo time steps per revolution set in S4: 1) For combinations containing rotor components, the final control matrix is ​​calculated based on the trim calculation parameters and trim target set in S5. The three-dimensional unsteady Reynolds-averaged NS (RANS) equations are solved based on the number of calculation revolutions, the physical time step per revolution, and the pseudo time step per revolution. The information of each component is exchanged through nesting relationships, and the inner background grid is exchanged with the outer background grid to finally obtain the disturbance flow field.

[0040] The formula for balancing is as follows:

[0041] The target quantity matrix of the rotor. This represents the rotor control matrix. (0< <1) is the damping factor. The Jacobian matrix calculated for balancing, This represents the difference in the target quantity matrix. This represents the difference in the manipulation matrix. The Jacobian matrix is ​​obtained through iterative calculation using the difference manipulation.

[0042] 2) For combinations containing propeller components, the three-dimensional unsteady Reynolds-averaged NS (RANS) equations are solved based on the number of revolutions, physical time steps per revolution, and pseudo time steps per revolution. Information of each component is exchanged through nesting relationships, and information is exchanged between the inner background mesh and the outer background mesh to finally obtain the disturbance flow field. 3) For a combination that includes at least one of the fuselage and wing components, the three-dimensional unsteady Reynolds-averaged NS (RANS) equations are solved based on the number of calculation revolutions, the physical time step per revolution, and the pseudo time step per revolution. The information of each component is exchanged through nesting relationships, and the inner background mesh is exchanged with the outer background mesh to finally obtain the disturbance flow field.

[0043] S65. Based on the disturbance flow field, obtain the aerodynamic force (six force elements), velocity field and vorticity field of each component under different physical time steps of each combination. To obtain the vorticity field at different physical time steps, see [link / reference]. Figure 5 ; To obtain the velocity field at different physical time steps, see [link / reference]. Figure 6 ; To obtain the unsteady aerodynamic forces of the rotor / fuselage / propeller in all aircraft configurations, see [reference]. Figure 7 ; To obtain the unsteady aerodynamic forces of components under different combinations, see [reference]. Figure 8 ; For unsteady aerodynamic harmonic analysis of the component, see [link / reference]. Figure 9 .

[0044] 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 unsteady aerodynamic disturbances in a multi-component aircraft, characterized in that, The method includes: S1, determine several combinations, each combination including at least one component: S2, perform mesh generation for each component; S3, divides all combinations into inner and outer background grids; S4 sets the calculation parameters for each combination. The calculation parameters include forward flight speed, density, number of calculation laps, physical time steps per lap, and pseudo time steps per lap. S5, for combinations that include rotor components, set the trim calculation parameters and trim targets; S6 calculates the disturbance flow field for each combination and obtains the unsteady disturbance law and mechanism of the component. The disturbance flow field includes aerodynamic force, velocity field and vorticity field.

2. The method according to claim 1, characterized in that, In S1, the combination with the most aircraft components is determined, and then other combinations are determined based on the characteristics of the aircraft components. Each combination includes at least one component; the components are divided into moving components and stationary components.

3. The method according to claim 1, characterized in that, In S2, for each component in S1, the static components are first locally trimmed, and the moving components are root-cut; then, each component is meshed, and the mesh type of each component is either a structured mesh or an unstructured mesh. For the moving component, the outer boundary is set to 1.5 to 1.8 times the section chord length, and the thickness of the first layer of the blade boundary layer mesh is set to the order of the corresponding chord length * 10^-5; the surface mesh of the blade leading edge and tip of the moving component is refined. For stationary components, set the outer boundary to 1.2 to 1.5 times the length of the reference stationary component, and set the thickness of the first layer of the boundary layer mesh of the stationary component to be on the order of the length of the reference stationary component * 10^-5; refine the surface mesh of the leading edge mesh of the stationary component.

4. The method according to claim 1, characterized in that, In S3, the inner background mesh is divided to include all components. The outer boundary of the inner background mesh is 1.5 to 2.0 times the maximum length of each component in each direction. The inner background mesh is densified in areas with complex flow of each component. The boundary of the outer background mesh in each direction is set to 5 to 10 times the length of the longest component. The mesh size ratio at the data exchange point between the outer and inner background meshes is 1 to 1.

2.

5. The method according to claim 1, characterized in that, In S4, For combinations that include rotor components, the calculation parameters also include the number of rotors, the total number of rotor blades, the rotor tip speed, the direction of rotation, and the rotor rotation center. For combinations that include propeller components, the calculation parameters also include the number of propellers, the total number of propeller blades, the propeller tip speed, the direction of propeller rotation, the propeller's axis of rotation, and the center of propeller rotation. For a combination that includes at least one of a fuselage and a wing component, the calculation parameters also include fuselage / wing attitude and fuselage / wing attitude reference position; Set up multi-level motion relationships for combinations containing moving parts; set up multi-level motion relationships for moving parts: first perform blade flapping motion and pitch change motion, then perform rotational motion; When realizing the flapping and pitch-changing motion of the propeller blades, a rotating coordinate system is established with the rotation center of the moving part as the origin. The X direction is the incoming flow direction, the Y direction is vertically downward, and the Z direction is determined by the right-hand rule. In the rotating coordinate system, the motion of the moving part considers the rotation, flapping, and pitch-changing motion of the rotor / propeller. The flapping and pitch-changing laws are expressed as functions of the azimuth angle. in, It is the blade azimuth angle. The coefficient of motion during the swing. The pitch motion coefficient; When achieving rotational motion, the rotational law is a function related to the tip velocity of the moving part.

6. The method according to claim 1, characterized in that, In S5, Set the trim calculation parameters, including: initial number of cycles, azimuth angle calculated in a single iteration, total number of trim iterations, trim target quantity, damping factor, and control difference calculated by Jacobi matrix; The trim targets set include four quantities: drag, lift, pitching moment, and rolling moment. Alternatively, the trim target can include three quantities: lift, pitch moment, and roll moment.

7. The method according to claim 1, characterized in that, S6 specifically refers to: S61. For each combination, read in the inner and outer background meshes and the calculation parameters; For combinations that include rotor components, it is also necessary to read in the rotor mesh, trim calculation parameters, and trim objectives; For assemblies that include propeller components, the propeller mesh also needs to be read in; For combinations that include fuselage / wing components, the fuselage / wing mesh also needs to be read in; S62. Perform hole-punching and contribution cell search on the inner background mesh and the outer background mesh, and store the nesting relationship between the two in an array; S63. Initialize the flow field information based on the forward velocity and density in the calculation parameters: assign the forward velocity and density variables to the grid divided in S2; For combinations that include rotor components, the forward velocity and density variables are assigned to the rotor mesh; For combinations that include propeller components, the forward velocity and density variables are assigned to the propeller mesh; For combinations that include fuselage / wing components, the forward velocity and density variables are assigned to the fuselage / wing mesh; S64. Calculate the disturbance flow field based on the number of calculation revolutions, physical time steps per revolution, and pseudo time steps per revolution set in S4: S65. Based on the disturbance flow field, obtain the aerodynamic force, velocity field, and vorticity field of the component under different physical time steps for each combination; compare and analyze the aerodynamic force, velocity field, and vorticity field of the component under different combinations to obtain the law and mechanism of unsteady disturbance to the component.

8. The method according to claim 7, characterized in that, S64 specifically refers to: For combinations containing rotor components, the final control matrix is ​​calculated based on the trim calculation parameters and trim objectives. The three-dimensional unsteady Reynolds-averaged NS (RANS) equations are solved based on the number of calculation revolutions, the physical time step per revolution, and the pseudo time step per revolution. Information of each component is exchanged through nesting relationships, and information is exchanged between the inner background mesh and the outer background mesh to finally obtain the disturbance flow field. The formula for balancing is: in, The target quantity matrix of the rotor. This represents the rotor control matrix. The damping factor is 0 < <1, The Jacobian matrix calculated for balancing, This represents the difference in the target quantity matrix. This indicates the manipulation matrix difference, and the Jacobian matrix is ​​obtained through iterative calculation of the difference manipulation. For a combination that includes a propeller component, or a combination that includes at least one of a fuselage and a wing component, the three-dimensional unsteady Reynolds-averaged NS (RANS) equations are solved based on the number of revolutions, the physical time step per revolution, and the pseudo time step per revolution. Information about each component is exchanged through nesting relationships, and information is exchanged between the inner background mesh and the outer background mesh, ultimately yielding the disturbance flow field.

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