A coupled aerodynamic simulation method for propeller-cabin-wing model

Abstract

This invention discloses a coupled aerodynamic simulation method for a rotor wing model, belonging to the field of computational fluid dynamics simulation and aircraft aerodynamic design technology. The method divides the computational domain into a non-slip region, a slip-rotation region, and a slip-stationary region, and separates the slip-rotation region and the slip-stationary region by a calculated optimal clearance region. Differentiated meshing strategies are applied to each region. Local refinement and optimization are implemented in the transition region, clearance region, and tip vortex separation region under the same position constraint. The optimal clearance distance for the rotor wing is determined by combining Courant number constraints and turbulence scale analysis. Finally, a pressure-based coupled solver and the SST kω turbulence model are used to complete the simulation. The simulation method of this invention significantly reduces numerical dissipation, avoids the accumulation of interpolation errors with iteration, significantly improves computational stability and accuracy, reduces server computing power requirements, and improves the aerodynamic simulation accuracy and computational stability of the coupled flow field.

Description

Technical Field

[0001] This invention belongs to the field of computational fluid dynamics simulation and aircraft aerodynamic design technology, specifically relating to a rotor wing model coupled aerodynamic simulation method. Background Technology

[0002] The integrated propeller-nacelle-wing configuration is a typical power arrangement for turboprop regional aircraft, tiltrotor aircraft, and other similar aircraft. In this configuration, the slipflow generated by the propeller rotation strongly interferes with the flow at the nacelle boundary layer and on the wing surface, resulting in complex flow phenomena such as tip vortex generation, gap leakage flow, and unsteady transmission at the rotating / stationary interface. Currently, numerical simulations of the coupled propeller-nacelle-wing aerodynamic problem mainly rely on the slip mesh method or the multi-reference frame method to handle the interaction between the rotating and stationary domains. However, several technical bottlenecks remain in achieving a balance between high accuracy and high efficiency:

[0003] (1) Conservation interpolation at the interface between slip and non-slip regions is difficult. Conventional slip grid interfaces often use interpolation algorithms for flux transfer. In regions with abrupt changes in grid size or large flow gradients, numerical dissipation and non-physical discontinuities are easily introduced, resulting in non-conservation of mass, momentum and energy fluxes, which directly reduces the prediction accuracy of unsteady thrust and torque pulsation.

[0004] (2) Mesh generation and flow capture in the gap region between the propeller hub and the nacelle are challenging. This gap is typically only 1% to 3% of the propeller diameter, with a small scale and drastic curvature changes. Using a uniform unstructured mesh makes it difficult to simultaneously consider near-wall y+ control, mesh orthogonality, and overall mesh size. If prism layers are simply inserted, the interlayer transition and polyhedral mesh connection are poor, easily generating distorted elements, leading to distortion of the pressure field and vortex structure within the gap, ultimately affecting the accurate calculation of nacelle drag and wing lift.

[0005] (3) Insufficient grid resolution in the propeller tip vortex separation zone. The propeller tip vortex has a significant impact on the downstream wing attachment line and separation zone. However, due to the grid refinement strategy, the propeller tip vortex often dissipates after transmitting several times the chord length due to numerical dissipation, and cannot truly reflect the interference between the wake and the wing. In particular, it has a weak ability to predict flow separation at non-design points.

[0006] (4) In complex geometries near the interface between rotating and stationary components, traditional mesh generation techniques cannot guarantee that the common boundary nodes are completely conformal. In particular, in the connection area between the rear end of the propeller hub and the nacelle, if a simple projection cutting method is used, gaps, overlaps and node misalignments often occur, leading to local flow anomalies and solution divergence.

[0007] The aforementioned problems make it difficult to achieve a balance between accuracy, stability, and engineering applicability in existing rotor-nacelle-wing coupled simulation methods. Therefore, there is an urgent need for a coupled aerodynamic simulation method that can comprehensively address interface conservation interpolation, high-resolution mesh generation for gaps, precise capture of tip vortices, and adaptive optimization of gap distances, thereby providing a reliable tool for engineering evaluation of integrated rotor-nacelle-wing designs. Summary of the Invention

[0008] This invention addresses the shortcomings of existing rotor-nacelle-wing model coupled aerodynamic simulation techniques by proposing a rotor-nacelle-wing model coupled aerodynamic simulation method, comprising the following steps:

[0009] S1: Import the simulation model. The computational domain of the simulation model includes a non-slip region and a slip region. The non-slip region and the slip region are three-dimensional spaces, and the slip region is located within the non-slip region. The non-slip regions include transition regions connected to the slip regions. The slip region includes a slip rotation region, a slip stationary region, and a gap region. The slip rotation region includes a propeller and its hub, and the slip rotation region is hemispherical. The slip stationary region includes a nacelle and a rigid connection between the nacelle and the wing. The gap region is located between the slip rotation region and the slip stationary region.

[0010] S2: Different regions are divided using different mesh types: non-slip regions use overlapping meshes; the sliding rotation region and the sliding stationary region use polyhedral meshes, and the gap region and near-wall region use prism layer meshes. The sliding rotation region, the sliding stationary region, and the gap region constitute the sliding region; the sliding mesh region and the non-slip region use interface meshes, with preset overlapping mesh interpolation and enabled flux conservation interpolation.

[0011] In step S1, the non-slip region adopts an overlapping mesh scheme. Compared with the traditional two-region scheme, which uses a cut-body mesh in the non-slip region, the number of meshes increases more gradually while ensuring equal orthogonality. This can effectively control the total number of meshes within the computing power capacity of ordinary servers.

[0012] In step S2, the sliding rotation region and the sliding stationary region adopt a polyhedral mesh. Compared with the traditional tetrahedral mesh, a single polyhedral unit contains more adjacent faces. For complex contours with large curvature changes, such as the propeller leading edge and blade root shaping, it can achieve better surface fitting with fewer units. Contour accuracy can be guaranteed without excessive local refinement, thus reducing mesh redundancy from the root. Under the same accuracy requirements, the total number of meshes is less, and the iterative convergence stability is higher.

[0013] In step S2, the gap region uses a prism layer mesh, including: setting the number of mesh layers to 5 to 8, setting the height of the first layer to 0.001m to 0.008m, and setting the height of the first layer to meet the control requirement of y+≈1 near the wall to ensure the calculation accuracy of the near-wall boundary layer flow field; setting the mesh growth rate to 1.1 to 1.3; generating prism layer and polyhedral mesh using shared nodes, with the outermost nodes of the prism layer and the nodes of the polyhedral mesh automatically matched and aligned without additional interpolation; the prism layer mesh and the polyhedral mesh together form a sliding mesh, and the overlapping parts between them are automatically removed.

[0014] The slip region further includes: a near-wall conformal region, a near-wall surface region, and a barrel-shaped blade tip vortex separation region; the near-wall conformal region is an annular area on the outer surface of the blade hub between the starting and ending rings, with the blade hub tip pointing forward as the forward direction and the connecting loop between the rearmost connection points of the blade hub and all blades as the starting loop and the outer ring of the blade hub bottom surface as the ending loop; the near-wall surface region is in close contact with the outer surface of the blade-nacelle-wing model other than the near-wall conformal region; the barrel-shaped blade tip vortex separation region is used to calculate the flow field parameters of the vortex generated by the simulated propeller rotation.

[0015] The near-wall region uses a prismatic layer mesh, the near-wall conformal region uses an imprinting process to generate a conformal mesh, and the barrel-shaped propeller tip vortex separation region uses a multi-layered, densely packed polyhedral mesh.

[0016] The near-wall conformal region is generated using an imprinting process to create a conformal mesh. Specific implementation steps include:

[0017] S21: Geometric model preprocessing, cleaning the model, repairing minor gaps and overlapping surfaces, ensuring that both boundaries are complete and undamaged closed surfaces, and avoiding imprint recognition failure;

[0018] S22: Mark the source boundary and target boundary. Select the sliding area end face on the moving side as the source boundary. Its contour is more regular and can be used as a high-precision imprint template. Select the adjacent end face of the non-rotating stationary area as the target boundary and receive the imprint contour. Keep the two boundaries parallel and set the gap to the optimized value.

[0019] S23: Perform imprinting projection, project the contour line of the source boundary onto the target boundary, cut out a new contour line on the target boundary that is completely consistent with the source contour, and divide the target boundary into an inner imprinting area and an outer region.

[0020] S24: Re-grid the target boundary. Based on the new contour obtained by imprinting, re-grid the planar mesh of the target boundary to ensure that the number of mesh nodes and the node distribution in the imprinted area correspond completely with the source boundary, and achieve dual alignment of contour and nodes.

[0021] S25: Generate a conformal mesh. Based on the aligned boundary mesh, stretch the generated volume mesh into the non-rotation region to ensure that the nodes of the two regions on the common boundary are completely coincident, and finally obtain the imprinted conformal mesh.

[0022] S26: Quality check and adjustment, check the node matching degree and mesh quality of the common boundary, remove unqualified distorted elements, and complete the final generation.

[0023] Step S26 includes: setting the average grid orthogonality of the common boundary region to be no less than 0.75, the aspect ratio to be no higher than 3:1, the aspect ratio at the common boundary to be no more than 5:1, and the gap region to be strictly controlled within 3:1; the contour matching accuracy to ensure that the deviation of the contour lines of the source boundary and the target boundary is no more than 10% of the minimum grid size; the node coincidence accuracy to ensure that the spatial position deviation of the corresponding nodes on the common boundary is no more than 5% of the minimum grid size; and the node number matching to ensure that the number of nodes of the source boundary and the imprinted target boundary are completely equal, and no redundant or missing nodes are allowed.

[0024] S3: Perform mesh optimization at the same location, and perform mesh adjustment and densification operations on the transition area and gap area between the slip area and the non-slip area, as well as the propeller blade tip and blade leading edge position.

[0025] The optimization at the same location in step S3 includes: fixing the global basic settings, keeping the computational domain size, far-field boundary conditions, turbulence model, Coulomb number constraints, propeller speed, and inlet / outlet conditions completely unchanged, and only adjusting the mesh partitions and gap sizes; fixing the total mesh size, controlling the deviation of the total mesh size before and after optimization to within ±5%, avoiding interference with the comparison results due to different mesh sizes; fixing the monitoring positions, setting surface thrust and torque monitoring points, SDR and TKE data sampling points, and eddy current monitoring sections on the outer layer of the 3D model, keeping them in the same physical location, and not changing the sampling positions due to partition adjustments to ensure data comparability; unifying the convergence judgment criteria, using the same residual convergence threshold before and after optimization, requiring a continuous residual decrease of 3 orders of magnitude, and a thrust monitoring value fluctuation of less than 0.5% to determine convergence, avoiding result deviations due to different convergence criteria; and saving the optimization results after multiple rounds of repeated verification.

[0026] The mesh adjustment and refinement operation for the transition region and gap region between the slip region and the non-slip region, as well as the key positions of the propeller blade tip and leading edge, includes:

[0027] The interstitial region is densified, and a local mesh size constraint is set separately for the optimized interstitial region. The minimum mesh size in the interstitial region is set to 1 / 10 to 1 / 8 of the interstitial width, and the expansion ratio is compressed to 1.1 to ensure the near-wall resolution of the prism layer in the interstitial region.

[0028] The transition adjustment between the sliding area and the background overlap area, and the transition adjustment between the sliding area and the non-sliding area, involves setting 3 to 5 layers of transition meshes between the sliding area and the non-sliding area. The mesh size gradually increases from the fine size of the sliding area to the coarse mesh size of the non-sliding area. The mesh growth rate of the transition mesh is set to within 1.2 to avoid numerical dissipation caused by abrupt changes in mesh size.

[0029] The critical locations of the propeller blade tip and the vortex separation region of the barrel tip are densified, and local densification is set separately for locations with large flow gradients; the outermost mesh and polyhedral mesh of the prism layer are optimized, and the difference between the size of the outermost mesh of the prism layer and the size of the adjacent polyhedral mesh is set to not exceed 2 times.

[0030] S4: The optimal clearance distance between propeller cells was determined by Courant number calculation, Taylor scale and Kolmogorov scale analysis, combined with case studies and server load capacity.

[0031] In step S4, the minimum gap distance is determined using the Courant number constraint, calculated as: CFL = U·Δt / Δx. To ensure numerical stability, the maximum Courant number within the gap must be less than or equal to 1, thus deducing that the minimum mesh size within the gap is Δx ≥ U·Δt. Since the gap needs to accommodate at least 8–10 mesh layers to resolve shear flow, the minimum gap requirement is calculated using the following formula:

[0032] G≥8·(U·Δt) and 1%D≤G≤3%D

[0033] Where G is the optimal clearance distance, U is the tangential velocity at the hub, Δt is the flow field time step, and D is the propeller diameter. The industry-standard optimal clearance value ranges from 1% to 3% of the propeller diameter. For propellers of general-purpose aircraft and UAVs, the optimal numerical range is 0.02m to 0.04m.

[0034] When the propeller clearance is insufficient, the mesh is prone to extreme stretching or negative volume, directly leading to computational divergence and collapse. Even if it does not collapse, poor mesh quality will cause strong numerical dissipation, resulting in an overestimation of propeller disturbance drag. When the propeller clearance exceeds the limit, the physical distance between the blade sweep flow field and the stationary hull will be exaggerated, leading to a weak simulation of the disturbance between the propeller slipstream and the hull, thus underestimating the total thrust loss and lift gain.

[0035] S5: Perform simulation calculations to obtain various aerodynamic performance indicators and flow field characteristic parameters.

[0036] The method uses STAR-CCM+ for mesh generation and simulation, employs a pressure-based coupled solver, and uses the SST k-ω model for turbulence. The SST k-ω model is chosen because it performs well in near-wall handling, predicting moderately separated flows, and capturing tip vortices, and is relatively robust to mesh quality. The time step corresponds to a propeller rotation of 1° to 2°, and the time step is calculated as: Δt = (1 / 360 to 1 / 180) × (60 / RPM) seconds.

[0037] Beneficial effects

[0038] This invention adopts a three-region architecture of "background overlapping area + double sliding area". The non-sliding area uses overlapping mesh, which avoids the problem that the number of cutting body meshes increases exponentially with the accuracy requirements. The sliding area uses polyhedral mesh, which is superior to traditional tetrahedral mesh in the reconstruction of complex contours.

[0039] While maintaining the same computational accuracy, the total mesh size of this scheme is much smaller than that of the traditional two-region cut-body mesh scheme, enabling high-precision simulations to be completed within the computing power range of ordinary servers. Imprinting a conformal mesh eliminates interpolation errors. A conformal mesh is obtained between non-rotated and non-slip regions through imprinting, with nodes completely overlapping on the common boundary, allowing for direct data exchange. Compared to traditional non-conformal AMI (Arbitrary Mesh Interface) interpolation, this significantly reduces numerical dissipation, avoids the accumulation of interpolation errors with iterations, and significantly improves computational stability and accuracy.

[0040] Quantitative calculations using the Courant number, Taylor scale, and Kolmogorov scale, combined with industry-standard parameters of 1%–3% of propeller diameter, and through multiple rounds of experimental verification, determined the optimal range for propeller clearance to be 0.02m–0.04m. This avoids the problems of computational divergence due to excessively small clearances or simulation distortion due to excessively large clearances. Furthermore, compared to the traditional two-region scheme, the three-region scheme of this invention exhibits more stable residual values, without any absurdly sharp increases or decreases in SDR, and more stable eddy current distribution and aerodynamic calculation results. It achieves long-term stable convergence and significantly reduces the probability of computational crashes due to mesh distortion. This invention completely solves the common industry problem in traditional non-conformal AMI interpolation schemes where errors accumulate with iterations, eventually leading to negative volume due to mesh distortion and server computational crashes.

[0041] This approach eliminates the inherent defect of the traditional two-region scheme where the propeller gap falls at the interface between the rotating and background regions from a physical layout perspective. It avoids the superposition effect of high shear gap flow and non-conformal interpolation errors, solves the problem of abnormally large SDR and TKE dissipation rates, improves simulation convergence stability, and significantly reduces numerical turbulence dissipation in the near-wall region. It also provides higher accuracy in calculating wall pressure distribution and wall shear stress, and can provide more reliable numerical simulation basis for coupled aerodynamic design of propeller-nacelle-wing model of aircraft turboprop engine. Attached Figure Description

[0042] Figure 1 Simulation scheme flowchart;

[0043] Figure 2 Three-region division and mesh generation effect diagram;

[0044] Figure 3 Optimization effect diagram of the near-wall region;

[0045] Figure 4 Transition area rendering;

[0046] Figure 5 Image showing the encrypted area;

[0047] Figure 6 Optimal gap area effect diagram;

[0048] Figure 7 Simulation results of the flow field in the three-region scheme;

[0049] Figure 8 Two-region division and mesh generation effect diagram;

[0050] Figure 9 Simulation results of the flow field in the two-region scheme. Detailed Implementation

[0051] To make the objectives, technical solutions, and advantages of this invention clearer, the technical solutions of this invention will be clearly and completely described below in conjunction with specific embodiments. Obviously, the described embodiments are only some, not all, of the embodiments of this invention.

[0052] Based on the embodiments of this invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this invention.

[0053] Example 1

[0054] This embodiment uses the rotor-nacelle-wing model coupled system of the JC-900 twin-engine turboprop aircraft as the simulation object, and employs the three-region coupled aerodynamic simulation method described in this invention for calculation. This embodiment selects the intermediate values ​​of each parameter range given in the technical solution for implementation. The specific steps are as follows:

[0055] 1. Simulation Model Preparation

[0056] A realistic geometric model of the nacelle was used, including a 2.4m diameter MT five-bladed propeller, the nacelle forebody, and the central wing connecting section. The propeller's rated speed is 1591 RPM, and the calculation condition used 96% of the rated speed, i.e., 1527 RPM. The diameter at the propeller hub root is 400mm. The geometric model was preprocessed to remove minor features with minimal impact on aerodynamic performance, repair tiny gaps and overlapping surfaces on the model surface, ensuring that all surfaces are intact and free of defects, and avoiding topological errors during subsequent mesh generation.

[0057] 2. Regional Division

[0058] Following the method described in step S1, the calculated flow field is divided into the following three functional regions:

[0059] Non-slip region: Covers the entire computational domain and is used to simulate the far-field flow field and the flow field in the rear region of the wing and nacelle. This region uses an overlapping mesh scheme.

[0060] Slip-rotating region: This region encompasses the entire propeller and hub. It is a hemispherical rotating region that encloses the rotating area of ​​the propeller and hub. It has a diameter of 2.6m and an axial length of 0.8m. The interior of this region uses a polyhedral grid to accurately reproduce the complex curved blade shape of the propeller.

[0061] The skid stationary region, comprising the nacelle and the rigid connection between the nacelle and the wing, is a cylindrical stationary region with a diameter of 2.6m, coaxially arranged with the skid rotation region. The axial clearance between the skid rotation region and the skid stationary region, based on the quantitative analysis method in step S4, is optimized to 0.032m (within the range of 0.02m to 0.04m, corresponding to approximately 1.33% of the propeller diameter of 2.4m). The skid rotation region, the skid stationary region, and the clearance region together constitute a capsule-shaped skid region, enclosing the overall propeller-nacelle-wing model.

[0062] 3. Hybrid Mesh Generation

[0063] Mesh generation was performed using STAR-CCM+ software, such as... Figures 2-4 As shown, each region executes the differentiated grid strategy described in step S2:

[0064] Non-slip region: An overlapping mesh is used with a base size of 0.8m. Local refinement is applied in the propeller / nacelle / wing interference area with a refinement size of 0.08m and a refinement factor of 10. This selection ensures good mesh orthogonality while avoiding the problem of exponential growth in mesh count near the wall in the cut-body mesh.

[0065] Slip region: The internal structure uses a polyhedral mesh with a basic size of 0.015m. The propeller leading edge and blade tip regions are key areas for mesh refinement, with a refinement size of 0.003m, to ensure sufficient capture of strong gradient flow fields and blade tip vortices. Compared to traditional tetrahedral meshes, polyhedral meshes offer higher surface fit in high curvature regions and require fewer mesh elements for the same level of accuracy.

[0066] Near-wall and gap regions: Prismatic meshes are generated on the surfaces of the propeller-cabin-wing model within the gap region between the sliding-rotating region and the sliding-stationary region, as well as within the sliding-rotating region and the sliding-stationary region. There are 5 prismatic layers, with the first layer having a height of 0.0035m (within the range of 0.001m to 0.008m). The mesh growth rate is 1.15 (within the range of 1.1 to 1.3). Calculations show that the near-wall y+ value meets the control requirement of y+≈1, ensuring the accuracy of the boundary layer flow field calculation. The prismatic layers and polyhedral meshes are generated using a shared-node method, requiring no additional interpolation.

[0067] Near-wall conformal region (imprinted conformal mesh): Perform imprinting processing according to steps S21 to S26. Perform geometric model preprocessing and clean up surface defects; then mark the end face of the moving side slip region as the source boundary and the adjacent end face of the non-rotating stationary region as the target boundary, keeping them parallel with a gap of 0.032m; perform imprinting projection to project the source boundary contour onto the target boundary and cut out a new contour line; re-mesh the target boundary plane mesh based on the new contour, ensuring that the number and distribution of nodes completely correspond to the source boundary; finally, stretch into the non-rotating region to generate a conformal mesh.

[0068] Quality testing revealed that the average mesh orthogonality in the common boundary region was 0.82, the aspect ratio was 2.3:1, the node coincidence accuracy met the requirement of a deviation not exceeding 4% of the minimum mesh size, and the number of nodes was perfectly matched. The tip vortex separation region employed a multi-layered, finely detailed polyhedral mesh to ensure full capture of tip vortex evolution.

[0069] The transition region between the slip and non-slip mesh regions uses an interface mesh, with pre-defined overlapping mesh interpolations and flux conservation interpolation enabled to ensure the conservation and transfer of mass, momentum, and energy between the rotating and stationary regions. The overlapping mesh automatically identifies and removes portions of the background mesh that overlap with the slip region, retaining only the effective computational cells.

[0070] 4. Mesh optimization

[0071] Perform mesh adjustment and refinement according to the same-location optimization method described in step S3:

[0072] (1) Gap region refinement: Local mesh size constraints are set separately for the 0.032m gap region, with the minimum mesh size in the gap being 3.2mm (i.e., 1 / 10 of the gap width). Preferably, at least 7 mesh layers are ensured in the gap, which is sufficient to capture the shear flow structure; the expansion ratio is compressed to 1.1 to fully ensure the near-wall resolution of the prism layer in the gap region, with the effect as follows: Figure 6 As shown.

[0073] (2) Adjustment of the transition between the slip zone and the background overlap zone: Four layers of transition mesh are set from the outer boundary of the slip zone to the background zone. The mesh size gradually increases from 0.015m in the slip zone to 0.8m in the background zone. The mesh growth rate is controlled at 1.18 (within 1.2) to avoid additional numerical dissipation caused by abrupt changes in mesh size.

[0074] (3) Refinement of critical locations at the propeller blade tip / leading edge: Local refinement layers are set separately for the propeller blade tip and leading edge regions to ensure at least 15 layers of mesh resolution in the tip vortex region, effectively capturing the vortex evolution process. The outermost mesh and polyhedral mesh of the prism layer are optimized, with the difference between the outermost prism layer mesh size and the adjacent polyhedral mesh size controlled within 1.8 times (not exceeding 2 times), ensuring smooth mesh transitions. The effect is as follows: Figure 5 As shown.

[0075] During the optimization process, the principle of comparison at the same location was strictly followed: all global parameters, including the computational domain size, far-field boundary conditions, turbulence model, Courant number constraints, propeller speed, and inlet / outlet conditions, were fixed, and only the grid partitioning and gap size were adjusted; the total number of grids before and after optimization was controlled within ±5%; the physical locations of all monitoring points were fixed; and a unified convergence criterion was used. After optimization, the total number of grids was over 8.6 million, which was kept within the computing power capacity of a typical server (16 cores and 32GB of memory).

[0076] 5. Solver Setup and Simulation Calculation

[0077] Perform simulation calculations according to the solution settings described in step S5. The specific parameter configurations are as follows:

[0078] Solver type: Pressure-based coupled solver, suitable for handling low-speed incompressible and subsonic compressible flow scenarios.

[0079] Turbulence Model: The SST k-ω two-equation turbulence model was adopted. The selection criteria were: (a) the model uses the k-ω formula near the wall, which can accurately analyze the viscous sublayer and meet the high accuracy requirements under the condition y+≤1; (b) it automatically switches to k-ε behavior in the far field, avoiding the sensitivity of the standard k-ω model to free flow; (c) it has excellent performance in predicting moderately separated flows and capturing tip vortices; (d) it is relatively robust to mesh quality and adapts to the complex mesh topology in the impeller gap region.

[0080] Time step: corresponding to a propeller rotation of 1.5°, calculated using the formula Δt=(1 / 360~1 / 180)×(60 / RPM), yields Δt≈1.64×10. - Four seconds, which is sufficient to accurately capture the transient rotational effect of the propeller.

[0081] Boundary condition settings: The far-field boundary is set to pressure far-field condition, the incoming flow velocity is 120m / s, the altitude is 0m, and standard atmospheric conditions are used (temperature 288.15K, pressure 101325Pa); the propeller region is set to rotational motion boundary with a speed of 1527RPM; the nacelle and wing surfaces are set to non-slip thermal insulation walls.

[0082] Convergence criteria: The calculation is considered converged when the continuous residuals of all solved variables decrease by more than three orders of magnitude and the fluctuation of the thrust monitoring value is less than 0.5%.

[0083] 6. Simulation Results

[0084] After convergence, the aerodynamic parameters of the JC-900 aircraft nacelle were obtained in two positions: the initial default position (in situ) and the position with a radial forward movement of 100 mm. The above simulation test procedure S1-S5 was repeated, and the aerodynamic performance parameters obtained are as follows:

[0085]

[0086] Combination Figure 9 Experimental data show that the maximum SDR value in the propeller gap region is approximately 8.7 × 10⁻⁶. 7 s -1 No sharp increases or decreases were observed during the entire calculation process, indicating that the flow field simulation effect in the gap region was good; the tip vortex structure was clear and complete, and there was no premature dissipation or breakup of the vortex due to numerical dissipation; the calculation process was stable, and there were no residual jumps or calculation crashes during 72 hours of continuous operation; the residuals of all monitored variables decreased steadily by more than 3 orders of magnitude, reaching the convergence criteria.

[0087] Example 2

[0088] This embodiment uses the same JC-900 aircraft rotor-nacelle-wing model coupling system as Embodiment 1 as the simulation object, with the geometric model and flight parameters remaining unchanged. This embodiment selects the left endpoint values ​​of each parameter range given in the technical solution for implementation, to verify that the method of the present invention can still operate effectively under the lower limit of the parameter range. Specific parameter values ​​are as follows:

[0089] 1. Zone division and gap setting

[0090] After completing the division of the three regions according to step S1, the optimal gap determination method in step S4 is used to select a gap distance of 0.02m.

[0091] 2. Hybrid Mesh Generation

[0092] The mesh type for both the non-slip and slip regions remains consistent with that in Example 1. Key parameters are all taken as the left endpoint values:

[0093] Prismatic layer mesh: The number of layers is set to 5, the height of the first layer is set to 0.0035m, and the mesh growth rate is set to 1.1. Calculations show that the y+ value near the wall still meets the control requirement of y+≈1.

[0094] Conformal region: The imprinting process is the same as in Example 1, with the gap set to 0.020m. Quality inspection showed that the average mesh orthogonality of the common boundary region was 0.78, the aspect ratio was 2.8:1, and the node coincidence accuracy met the requirement that the deviation should not exceed 5% of the minimum mesh size.

[0095] 3. Mesh optimization

[0096] Densification of the gap area: The minimum mesh size within the gap is 1 / 8 of the gap width, i.e., 2.5mm, ensuring at least 8 mesh layers within the gap. The expansion ratio is compressed to 1.1.

[0097] Adjustment of the transition between the sliding area and the background overlap area: Set 3 layers of transition mesh, and set the mesh growth rate to 1.2.

[0098] Refinement of critical locations at propeller blade tips / leading edges: Consistent with Example 1, the difference between the outermost grid size of the prism layer and the adjacent polyhedral grid size is controlled within 2 times.

[0099] 4. Solver Settings

[0100] Time step Δt: Set as the time it takes for the propeller to rotate 1°, according to the formula Δt=(1 / 360)×(60 / 1527)≈1.09×10 -4 Seconds. All other solver parameters remain the same as in Example 1.

[0101] 5. Simulation Results

[0102] After optimization, the total grid size is approximately 7.5 million, which is within the computing power range of a typical server with 16 cores and 32GB of memory. The computation process is stable, with residuals consistently decreasing by more than three orders of magnitude without any crashes or divergences. The residuals of all monitored variables meet the convergence criteria. The flow field simulation in the gap region is effective, with no abnormal fluctuations in SDR and TKE, and the tip vortex structure remains intact. Simulation results demonstrate that, under the condition that all parameters are taken at the left endpoint of the range given in the technical solution, the method of this invention can still run stably and obtain reliable aerodynamic simulation results.

[0103] Example 3

[0104] This embodiment uses the same JC-900 aircraft rotor-nacelle-wing model coupling system as Embodiment 1 as the simulation object, with the geometric model and flight condition parameters remaining unchanged. This embodiment selects the right-end values ​​of each parameter range given in the technical solution for implementation, to verify that the method of the present invention can still operate effectively under the upper limit of the parameter range. Specific parameter values ​​are as follows:

[0105] 1. Zone division and gap setting

[0106] After completing the division of the three regions according to step S1, the gap distance is selected as 0.04m according to the gap determination method in step S4.

[0107] 2. Hybrid Mesh Generation

[0108] The mesh type for both the non-slip and slip regions remains consistent with that in Example 1. Key parameters are all taken as the right endpoint values:

[0109] Prismatic layer mesh: The number of layers is set to 8, the height of the first layer is set to 0.0035m, and the mesh growth rate is set to 1.2. Calculations show that due to the increased height of the first layer, the y+ value near the wall increases slightly but still meets the control requirement of y+≈1.

[0110] Conformal region: The imprinting process is the same as in Example 1, with the gap set to 0.04m. Quality inspection showed that the average mesh orthogonality of the common boundary region was 0.80, the aspect ratio was 2.5:1, and the node coincidence accuracy met the requirement that the deviation should not exceed 5% of the minimum mesh size.

[0111] 3. Mesh optimization

[0112] Densification of the gap area: The minimum mesh size within the gap is 1 / 10 of the gap width, i.e., 4.0 mm, ensuring at least 10 mesh layers within the gap. The expansion ratio is compressed to 1.1.

[0113] Adjustment of the transition between the sliding area and the background overlap area: Set 5 layers of transition mesh, and set the mesh growth rate to 1.2.

[0114] Refinement of critical locations at propeller blade tips / leading edges: Consistent with Example 1, the difference between the outermost grid size of the prism layer and the adjacent polyhedral grid size is controlled within 2 times.

[0115] 4. Solver Settings

[0116] Time step: Set to the time it takes for the propeller to rotate 2°, according to the formula Δt=(1 / 180)×(60 / 1527)≈2.18×10 -4 Seconds. All other solver parameters remain the same as in Example 1.

[0117] 5. Simulation Results

[0118] After optimization, the total grid size is approximately 9.5 million, which is within the computing power range of a typical server with 16 cores and 32GB of memory. The computation process is stable, with residuals consistently decreasing by more than three orders of magnitude without any crashes or divergences. The residuals of all monitored variables meet the convergence criteria. The flow field simulation in the gap region is effective, with no abnormal fluctuations in SDR and TKE, and the tip vortex structure remains intact. Simulation results demonstrate that, under the condition that all parameters are taken at the right-end values ​​of the given range of the technical solution, the method of this invention can still run stably and obtain reliable aerodynamic simulation results.

[0119] Comparative Example: Two-Region Scheme

[0120] To illustrate the technological advancements of this invention, comparative experiments were conducted under the same operating conditions, such as... Figure 8 As shown, the comparative example two region scheme adopts the architecture of "single sliding rotation region + single background region", the propeller gap is set to 0.010m, and the background region adopts a cut volume mesh.

[0121] The comparative simulation results show that: 1. The residual curve exhibits obvious periodic oscillations and cannot converge stably to below 3 orders of magnitude; 2. The SDR and TKE at the propeller gap exhibit sharp-peaked abnormal fluctuations, and the spatial distribution of turbulence parameters deviates systematically from the physical expectations; 3. The mesh in the gap region is distorted and produces negative volume, causing the solver to crash and failing to complete the complete simulation process.

[0122] pass Figure 7 and Figure 9 The comparative analysis of the experimental results shows that the above comparison fully verifies the significant improvement of the three-region scheme of the present invention in terms of computational stability, numerical accuracy and engineering practicality compared with the traditional two-region scheme; and the parameter range verification of Examples 1 to 3 further proves that the parameter ranges given by the technical solution of the present invention are feasible and effective throughout the entire interval.

[0123] The above description is merely a preferred embodiment of the present invention and is not intended to limit the present invention in any way. Although the present invention has been disclosed above with reference to preferred embodiments, it is not intended to limit the present invention. Any person skilled in the art can make some modifications or alterations to the above-disclosed technical content to create equivalent embodiments without departing from the scope of the present invention. Any simple modifications, equivalent changes and alterations made to the above embodiments based on the technical essence of the present invention without departing from the scope of the present invention shall still fall within the scope of the present invention.

Claims

1. A coupled aerodynamic simulation method for a propeller-cabin-wing model, comprising the following steps: S1: Import the simulation model, the computational domain of which includes the non-slip region and the slip region; The non-slip region and the slip region are in three-dimensional space, and the slip region is located within the non-slip region; The non-slip regions include transition regions that connect with the slip regions; The sliding region includes a sliding rotation region, a sliding stationary region, and a gap region; The slip rotation region includes the propeller and its hub, and the slip rotation region is hemispherical; The skid stationary zone includes the nacelle and the rigid connection between the nacelle and the wing; The gap region is provided between the sliding rotation region and the sliding stationary region; S2: Different regions are divided using different mesh types. The non-slip region uses an overlapping mesh; the slip-rotation region and the slip-stationary region use a polyhedral mesh; the initial gap region uses a prism layer mesh; the transition region between the slip region and the non-slip region uses an interface mesh. An overlapping mesh interpolation is preset in the interface mesh, and flux conservation interpolation is enabled. S3: Perform on-site mesh optimization on the mesh of the computational domain region, and perform mesh adjustment and refinement operations on the transition region between the slip region and the non-slip region, the gap region, and the propeller blade tip and leading edge position; S4: By calculating the Courant number, analyzing the Taylor and Kolmogorov scales, and combining numerical examples and server load capacity, the optimal clearance distance between the propeller cells is determined, and the optimized computational domain is obtained. S5: Perform simulation calculations on the optimized computational domain to obtain the aerodynamic performance indicators and flow field characteristic parameters of the propeller-nacelle-wing model within the computational domain.

2. The propeller-cabin-wing model coupled aerodynamic simulation method according to claim 1, characterized in that, The gap region uses a prism layer mesh, including: setting the number of mesh layers to 5 to 8, setting the height of the first layer to 0.001m to 0.008m; setting the mesh growth rate to 1.1 to 1.3; the prism layer and the polyhedral mesh are generated using shared nodes, and the outermost nodes of the prism layer and the nodes of the polyhedral mesh are automatically matched and aligned without the need for additional interpolation.

3. The propeller-cabin-wing model coupled aerodynamic simulation method according to claim 1, characterized in that, The sliding region also includes: a near-wall conformal region, a near-wall surface region, and a barrel-shaped blade tip vortex separation region; With the direction of the hub tip as the forward direction, the loop connecting the hub and the rearmost connection point of all blades is the starting loop, and the outer ring of the bottom surface of the hub is the ending loop. The annular area on the outer surface of the hub between the starting loop and the ending loop is the near-wall conformal region. The near-wall region is in close contact with the outer surface of the propeller-cabin-wing model, excluding the near-wall conformal region. The barrel-shaped propeller tip vortex separation region is used to calculate the flow field parameters of the vortex generated by the simulated propeller rotation.

4. The propeller-cabin-wing model coupled aerodynamic simulation method according to claim 3, characterized in that, The near-wall region uses a prismatic layer mesh, and the near-wall conformal region uses an imprinting process to generate a conformal mesh; the barrel-shaped propeller tip vortex separation region uses a multi-layered, densely packed polyhedral mesh.

5. The propeller-cabin-wing model coupled aerodynamic simulation method according to claim 4, characterized in that, in, The near-wall conformal region is generated into a conformal mesh using an imprinting process. Specific implementation steps include: S21: Geometric model preprocessing, cleaning the model, repairing minor gaps and overlapping surfaces, ensuring that both boundaries are complete and undamaged closed surfaces, and avoiding imprint recognition failure; S22: Mark the source boundary and target boundary. Select the sliding area end face on the moving side as the source boundary. Its contour is more regular and can be used as a high-precision imprint template. Select the adjacent end face of the non-rotating stationary area as the target boundary and receive the imprint contour. Keep the two boundaries parallel and set the gap to the optimized value. S23: Perform imprinting projection, project the contour line of the source boundary onto the target boundary, cut out a new contour line on the target boundary that is completely consistent with the source contour, and divide the target boundary into an inner imprinting area and an outer region. S24: Re-grid the target boundary. Based on the new contour obtained by imprinting, re-grid the planar mesh of the target boundary to ensure that the number of mesh nodes and the node distribution in the imprinted area correspond completely with the source boundary, and achieve dual alignment of contour and nodes. S25: Generate a conformal mesh. Based on the aligned boundary mesh, stretch the generated volume mesh into the non-rotation region to ensure that the nodes of the two regions on the common boundary are completely coincident, and finally obtain the imprinted conformal mesh. S26: Quality check and adjustment, check the node matching degree and mesh quality of the common boundary, remove unqualified distorted elements, and complete the final generation.

6. The propeller-cabin-wing model coupled aerodynamic simulation method according to claim 5, characterized in that, Step S26 further includes: The average mesh orthogonality of the common boundary region is set to be no less than 0.75, the aspect ratio is no higher than 3:1, the aspect ratio at the common boundary does not exceed 5:1, and the gap region is strictly controlled within 3:1; the contour matching accuracy meets the requirement that the deviation of the contour lines of the source boundary and the target boundary is no greater than 10% of the minimum mesh size; the node coincidence accuracy meets the requirement that the spatial position deviation of the corresponding nodes on the common boundary is no greater than 5% of the minimum mesh size; the node number matching meets the requirement that the number of nodes of the source boundary and the imprinted target boundary are completely equal, and no redundant or missing nodes are allowed.

7. The propeller-cabin-wing model coupled aerodynamic simulation method according to claim 1, characterized in that, The same-position optimization includes, Fix the global basic settings and adjust the mesh size of each region. Fixing the global basic settings includes keeping the computational domain size, far-field boundary conditions, turbulence model, Coulomb number constraints, propeller speed and inlet / outlet conditions unchanged. The total mesh size was kept constant, and the deviation of the total mesh size before and after optimization was controlled within ±5%. The monitoring positions were fixed, and the surface thrust and torque monitoring points, Specific Dissipation Rate (SDR) and Turbulent Kinetic Energy (TKE) data sampling points and eddy current monitoring sections were all set at the same physical location on the outer layer of the 3D model. The convergence criteria were unified, and the same residual convergence threshold was used before and after optimization. Convergence was judged when the continuous residual decreased by 3 orders of magnitude and the thrust monitoring value fluctuation was less than 0.5%.

8. The aerodynamic simulation method for a coupled propeller-cabin-wing model according to claim 1, characterized in that, The mesh adjustment and refinement operations are performed on the transition region, the gap region, and the propeller blade tip and leading edge positions, specifically including: The interstitial region is densified, and a local mesh size constraint is set separately for the optimized interstitial region. The minimum mesh size in the interstitial region is set to 1 / 10 to 1 / 8 of the interstitial width, and the expansion ratio is compressed to 1.1 to ensure the near-wall resolution of the prism layer in the interstitial region. The transition adjustment between the slip zone and the non-slip zone involves setting 3 to 5 layers of transition grids within the transition zone. The grid size gradually increases from the fine size of the slip zone to the coarse size of the non-slip zone, and the grid growth rate of the transition grid is set to within 1.

2. The critical locations of the propeller blade tip and tip vortex separation region are densified, and local densification is set separately for locations with large flow gradients; the prism layer mesh and polyhedral mesh are optimized, and the difference between the outermost mesh size of the prism layer mesh and the adjacent polyhedral mesh size is set to not exceed 2 times.

9. The propeller-cabin-wing model coupled aerodynamic simulation method according to claim 1, characterized in that, Step S4, determining the optimal clearance distance between propeller cells, includes: Calculate the minimum mesh size within the gap region while ensuring that the maximum coulomb size within the gap region is less than or equal to 1; set the gap region to accommodate at least 8-10 mesh layers, and the formula for calculating the optimal gap distance between propeller modules is: G≥8·(U·Δt) and 1%D≤G≤3%D Where G is the optimal clearance distance, U is the tangential velocity at the hub, Δt is the flow field time step, and D is the propeller diameter.

10. The propeller-cabin-wing model coupled aerodynamic simulation method according to claim 1, characterized in that, The simulation calculation includes mesh generation and simulation solution. The solver adopts a pressure-based coupled solver, and the turbulence model adopts the SST k-ω model. The time step is set to the duration of propeller rotation of 1° to 2°.

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