22 results about "Blade geometry" patented technology

The invention discloses a multidisciplinary coupling adjoint optimization method for a large-bypass-ratio fan blade. The method comprises the following steps of: 1, extracting two-dimensional blade profiles of different blade heights of a large-bypass-ratio fan blade and blade modeling parameters of each blade height; 2, obtaining optimization design parameters; 3, obtaining an unsteady harmonic balance equation according to the unsteady flow field equation, and calculating a pneumatic-aeroelastic multidisciplinary coupling objective function; 4, establishing an unsteady adjoint equation, and performing iterative solution to obtain an unsteady adjoint variable; 5, obtaining sensitivity information of the aerodynamic-aeroelastic multidisciplinary coupling objective function about design parameters; step 6, updating and optimizing design parameters; 7, updating blade modeling parameters and blade geometry, and updating the computational grid by adopting a linear elasticity method; and 8, when the two norms of the sensitivity information meet the convergence standard, outputting the optimized blade modeling parameters and blade geometry. The method has the advantages of being low in calculation cost, high in optimization design automation degree and high in optimization efficiency.

Techniques for conducting an air flow simulation for a wind turbine include importing a file containing a digitized representation of a three-dimensional blade geometry, extracting from the file, blade constructive parameters, and calculating a low-order air flow past a wind-turbine that includes the blade, based on a Blade Element Momentum Theory (BEMT) to determine sectional angle of attack and free-stream velocity, boundary layer transition, and acoustic noise results. The techniques also include performing air flow simulation for a given number of blade sections, and generating virtual microphone rings. The process also includes computing noise spectra at the virtual microphone rings and blending the noise spectra generated and generating synthetic noise signals from each section by inverse Fourier transform of the noise spectra and converting the noise spectra into an audio track.

The invention described herein pertains to a housing for a vertical-axis wind turbine, a vertical-axis wind turbine, and an apparatus comprising a housing and a vertical-axis wind turbine. For example, the disclosure below provides more efficient electricity generation from a vertical-axis wind turbine assembly. Increases in efficiency may be achieved by the housing through novel angled louvres and integrally housed electrical stators. Increases in efficiency may also be achieved through a magnetic spindle bearing and Savonius blade geometry. The turbine may be sized and configured for attachment to existing structures.

The invention provides an active control aeroelastic modeling method and device for a flap of a half-unfolded tilt-rotor aircraft, and the method comprises the steps: integrating a flap model into a cantilever wing model through a trailing edge flap interface and flap structure parameters reserved at a rotor structure assembly through the dynamic comprehensive analysis software of the rotor aircraft, establishing a cantilever wing model with flap active control; establishing a tilt rotor model; the cantilever wing model with the flap active control function and the tilt rotor model are combined together, and a flap active control model of the half-unfolded tilt rotor aircraft is obtained; carrying out aeroelastic stability initialization calculation on the flap active control model of the half-spread tilt-rotor aircraft, and extracting geometric shape data of a flap active control rotor blade of the half-spread tilt-rotor aircraft, so that the flap active control model can be integrated on a wing of a tilt rotor; the semi-expanded tilt rotor active control scaling model is created, and theoretical support is provided for follow-up tilt rotor flap active control aeroelastic design.

The application belongs to the technical field of turbine blade design. The application provides a turbine blade parameterized modeling method and system based on integrated design of blade end wall. The complete three-dimensional turbine blade geometric surface is directly generated by NFFD grid in the embodiment of the disclosure, so that the high-order smooth transition between the blade and the end wall is easy to realize. The method not only supports local modification of the blade geometry, but also provides rich geometric modification deformation parameters, which facilitates determination of the adjustment range of the parameters and ensures that unreasonable geometric shapes do not occur. In the optimization design process, no additional constraint conditions need to be set, so that a flexible and efficient optimization process is realized. In addition, the method is suitable for modeling and parameterization of axial turbine blades for various purposes, and has good regulation and control effect on the end region secondary flow.

The present disclosure provides a method and device for determining a blade deformation simulation parameter and an electronic device. The method comprises: generating a deformation driving model of a blade model based on blade modeling information, the deformation driving model comprising a plurality of sub-models, each sub-model being used to drive the blade model to produce deformation in an activated state; optimizing activation information of the deformation driving model using a Bayesian optimization algorithm to determine target activation information of the deformation driving model corresponding to a target blade geometry; and determining the target activation information of the deformation driving model corresponding to the target blade geometry as the blade deformation simulation parameter. The method provided by the present disclosure can not only realize forward design of the blade producing target deformation, but also drive the blade to actively produce deformation under different engine operating conditions by designing target activation information of different blade geometries to obtain corresponding blade geometries, so that the engine can have good aerodynamic performance in a wide speed range.

This invention discloses a compressor sand ingestion wear robustness optimization design method, system, equipment, and medium. First, preliminary aerodynamic optimization is performed to ensure that the compressor's factory performance does not degrade compared to traditional designs. Then, a high-precision real-time simulation method for compressor blade wear is constructed. By updating the coupling relationship between blade geometry and wear rate in real time, it accurately simulates blade geometric wear and aerodynamic performance degradation caused by sand ingestion during compressor operation in an outdoor environment. Next, based on this real-time simulation method, sand ingestion uncertainty analysis is performed on the preliminary design scheme, and wear-sensitive design parameters are optimized to improve the compressor's sand ingestion wear robustness, resulting in an optimized design scheme. Finally, the aerodynamic performance of the optimized design scheme is verified. This method can improve the wear resistance of compressors without reducing their aerodynamic performance as in traditional designs. The entire optimization process requires no manual design experience and helps alleviate the performance degradation of turboshaft engines.

This invention relates to the field of compressor aerodynamic performance evaluation technology, and discloses a method for quantitatively evaluating the uncertainty of blade geometric deviation performance. The method includes: acquiring the original design geometry of a multi-stage compressor and performing CFD simulation to determine the target reference flow rate; extracting deviation parameters to generate a deviation geometry sample set; performing multi-row CFD simulations on the samples; dynamically adjusting the outlet back pressure to constrain each sample to converge under the target reference flow rate; acquiring aerodynamic response data; training a surrogate model using the deviation parameters and aerodynamic response data; using the surrogate model for random sampling inference to obtain the probability distribution characteristics of the aerodynamic response data; combining a sensitivity analysis algorithm to calculate the global feature importance of each geometric deviation parameter and outputting core sensitive parameters; and determining the aerodynamic performance boundary and stall risk threshold of the compressor blade under operating conditions. This invention solves the problem of simulation condition deviation, reduces the computational cost of evaluation, and provides an engineering quantitative basis for stall risk early warning.

This application provides a dual-design-point calculation method for a compression system based on a one-dimensional inverse-forward-inverse problem, belonging to the field of gas turbine engine technology. The method includes obtaining the boundary conditions and blade geometry parameters of the first design point; performing a one-dimensional inverse problem analysis to obtain the one-dimensional design parameters of the first design point and transferring them to the second design point; obtaining the compression system's speed, flow rate, and pressure ratio at the second design point; performing a one-dimensional forward problem analysis to obtain the current flow rate at the second design point; performing another one-dimensional inverse problem analysis; calculating whether the pressure ratio and efficiency at the second design point meet the requirements based on the current flow rate; if not, adjusting according to a step-by-step process until the pressure ratio and efficiency at the second design point meet the requirements; and calculating the one-dimensional design parameters of the compression system at the second design point to obtain the one-dimensional key performance influencing parameters at both the first and second design points. This application improves the accuracy and efficiency of dual-design-point calculation for compression systems, and enhances its versatility and flexibility.

The invention belongs to the technical field of aero-engine manufacturing and maintenance, and particularly discloses an aero-engine regional damaged blade geometric modeling method which comprises the following steps: S1, determining the position of a blade profile surface damaged region and corresponding blade thickness and chord length damage parameters, and combining point cloud data of each section in a blade design model to obtain a blade surface damaged region; positioning the cross section of the damaged area; s2, performing offset processing on the point cloud data of each cross section of the damaged area; s3, importing the processed damage area point cloud data and non-damage area point cloud data into modeling software, adjusting geometric parameters of the blade edge to enable the non-damage edge position of the blade damage model to be consistent with the blade design model, and enabling the section chord length of the blade damage model to accord with a preset damage feature size; and S4, geometric modeling. According to the method, the blade geometric model with different degrees of regional damage characteristics can be efficiently constructed so as to support aerodynamic performance comparative analysis in different damage states.

The present application belongs to the technical field of aero-engine high-temperature component test, and particularly relates to a kind of turbine blade rotating environment under thermo-mechanical complex fatigue test equipment, including high-temperature gas system, stator system, rotor system, support system, test system, auxiliary system, turbine rotor blade test piece, the high-temperature gas system is used to provide gas environment for turbine rotor blade test piece, and the high-low temperature cycle state of gas is controlled, its front end is equipped with combustion chamber, combustion chamber is equipped with double oil path nozzle, rear end is equipped with gas flow system, gas flow system is designed cascade structure according to turbine rotor blade geometry, and is equipped with spray cooling section in test section rear part.The present application is composed of six core systems of high-temperature gas system, stator system, rotor system, support system, test system and auxiliary system, each system works cooperatively, realizes the integration simulation of "high-temperature gas environment + rotating centrifugal load + heat-power synchronous cycle".

The present application relates to a kind of end wall circumferential monotone blade end wall fusion design method, belong to aero-engine design field.First according to maximum fusion thickness, position and front trailing edge fusion scale determine blade fusion shape, then according to the flow characteristics of transverse secondary flow determine end wall monotone slope angle and position, finally according to the constraint of each axial position passage area invariable, determine end wall and blade geometry, realize end wall circumferential monotone blade end wall fusion design.This method strengthens the control ability of the angle area dihedral angle distribution formed by end wall and blade, can control the fusion shape of suction surface and leading edge, trailing edge simultaneously, root blade element section and end wall are included in aerodynamic design space, ensure that the passage area of each axial section is invariable, avoid the mass flow variation caused by metal blockage area change and working condition deviation, more have engineering application value.End wall modeling uses circumferential monotone curve structure, while controlling transverse secondary flow, ensure good working condition adaptability.

The invention provides a vibration and noise reduction blade design method based on flow channel and load cooperative regulation and control, an impeller and a vane pump, and the method comprises the steps: defining blade meridian plane flow channel shape determination parameters, including a rim inlet diameter, a hub inlet diameter, a rim flow channel curvature radius, a hub flow channel curvature radius, a rim outlet inclination angle and a hub outlet inclination angle; the geometric shape of the blade is determined according to the blade loads of the hub, the rim and a plurality of molded lines located between the hub and the rim at different blade heights. One purpose of the invention is to provide a vibration and noise reduction blade design method based on cooperative regulation and control of the flow channel and the load, and according to the method, the blade load distribution at different blade height positions is corrected by utilizing the curvature of the flow channel on the meridian plane of the blade, so that unsteady and non-uniform fluid excitation in an impeller can be effectively inhibited; the method has important significance on vibration and noise reduction of hydraulic machinery. The impeller is easy to implement and suitable for various types of impellers.

This application discloses a rotor aerodynamic structure coupling analysis method, apparatus, equipment, and medium, relating to the field of rotor performance analysis. The method includes: inputting rotor operating parameters and model parameters at the current aerodynamic time step into rotor blade models of different fidelity for solution, obtaining current aerodynamic load parameters corresponding to different fidelity levels; inputting the current rotor operating parameters and aerodynamic load parameters into a structural deformation prediction model to obtain the current rotor structural deformation; the structural deformation prediction model is determined based on a reduced-order modeling method and a neural network model; updating the spatial positions of each section of the blade according to the current rotor structural deformation at different fidelity levels, changing the blade geometry and aerodynamic boundary conditions, and using the changed geometry and aerodynamic boundary conditions as model parameters for the next aerodynamic time step, until the rotor structural deformation at all aerodynamic time steps is obtained. This application can reduce overall computational costs while ensuring aerodynamic analysis accuracy.

The invention relates to the technical field of wind driven generator blade layering, in particular to a reverse design method, device and equipment for blade layering and a computer readable storage medium. The method comprises the steps that target bending rigidity distribution of a blade and the geometric shape of the blade are obtained, and a basic shell layering is laid on the geometric shape of the blade; a beam cap is arranged on the laying layer of the foundation shell with the blade shaft as the center line; aiming at the blade section, calculating the initial bending rigidity of a foundation shell laying layer when no beam cap exists; and judging whether the target bending rigidity is greater than the initial bending rigidity or not, if so, sequentially increasing the number of unidirectional laying layers from the first beam cap, and calculating the bending rigidity of the current interface until the error between the bending rigidity of the current section and the target bending rigidity meets a preset tolerance. Target bending rigidity distribution is used as direct input, laying layer parameters are reversely solved through a systematized algorithm process, and a large amount of trial and error calculation and repeated adjustment depending on experience in a traditional method are avoided.

The invention discloses a CFD (computational fluid dynamics) method-based wide-speed-range efficient ducted fan aerodynamic design method and a CFD method-based wide-speed-range efficient ducted fan aerodynamic design device. According to the method, parametric modeling is conducted on a ducted fan, a CFD calculation model suitable for various working conditions such as hovering and cruising is constructed, hovering efficiency and cruising propulsion efficiency serve as design targets, a multi-target optimization function is called, CFD simulation iteration is driven through an optimization algorithm, and a Pareto optimal solution set is obtained and serves as a candidate scheme. And then a proper scheme is selected from the solution set according to actual requirements, local correction is carried out through flow field diagnosis, and finally a globally optimized ducted fan geometric model is obtained. According to the ducted fan device designed according to the method, the ducted molded surface and the blade geometry are optimized, the efficient aerodynamic configuration is achieved, the performance contradiction in the hovering state and the cruising state can be effectively coordinated, the high aerodynamic efficiency is achieved in the wide speed domain range, and the ducted fan device is particularly suitable for propelling systems of aircrafts such as eVTOL and the like.

This utility model relates to the field of multi-blade centrifugal fan technology, specifically to an impeller structure, a multi-blade centrifugal fan, and a multi-blade centrifugal blower. It includes a first disc and a second disc, and several blades located between them. The blades have an arc-shaped cross-section and are spirally twisted from the end closer to the first disc to the end closer to the second disc. A first centerline L1 exists at the end of the blade closer to the first disc, and a second centerline L2 exists at the end of the blade closer to the second disc. The projections of the first centerline L1 and the second centerline L2 intersect at an intersection point P. The intersection point P, the midpoint m1 of the first centerline L1, and the midpoint m2 of the second centerline L2 do not coincide simultaneously. This utility model optimizes the blade geometry, effectively improving the airflow guidance capability, improving airflow organization, and significantly reducing eddies and flow separation phenomena, thereby improving the aerodynamic performance of the blower. Furthermore, it improves the impeller's operating efficiency while effectively reducing operating noise.

An aircraft propulsion system includes a bladed rotor configured for rotation about a rotational axis. A rotor blade of the bladed rotor of rotor blades includes a blade body. The blade body has a blade span extending between and to a base end and a tip end. The blade body has a blade chord extending between and to a leading edge and a trailing edge. The blade body forms a pressure side surface and a suction side surface. The blade body forms a recess at the pressure side surface. The recess includes a recess surface portion forming a portion of the pressure side surface. The recess surface portion has a recess height. The recess height is greater than or equal to a Y-value of 0.025 for a normalized two-dimensional curve of a surface profile of the recess surface portion taken along a Y-Z plane orthogonal to the rotational axis for a recess span.

The invention relates to the technical field of fluid machinery design, in particular to an impeller design method based on cylindrical projection, which comprises the following steps: firstly, three-dimensionally and geometrically decomposing an impeller into a layered structure of a hub layer, a central layer and a wheel cover layer, and representing the contour of each layer by adopting a three-dimensional space curve; projecting and equally dividing the space curve of each layer by taking the cylindrical surface of the outlet of the impeller as a projection reference surface; secondly, initial geometric parameters are determined according to actual working condition requirements, and geometric parameters of all stepping points are dynamically solved through collaborative iterative calculation of an axial molded line function and a radial molded line function; and finally, fitting to generate an optimized three-dimensional contour of the impeller. According to the method, through parameterized layered modeling and explicit function correlation, accurate matching of blade geometry and flow field characteristics is achieved, the problem of flow field distortion in traditional design is effectively solved, spanwise profile control is remarkably improved, secondary flow and separation loss are reduced, and the design efficiency is guaranteed while the design cost is reduced. And the aerodynamic performance and the structural reliability of the impeller under severe working conditions are greatly improved.

The invention provides a vibration and noise reduction blade design method based on quadric function airfoil thickness distribution, an impeller and a blade. The vibration and noise reduction blade design method comprises the steps that a performance target is determined; determining physical boundary conditions of the blade; a thickness distribution control law of reverse solving is determined, a mathematical equation set with a quartic polynomial function of the blade thickness changing along the blade height as a solving target is constructed, and coefficients are undetermined unknown numbers; a two-dimensional airfoil section is determined, a parameterized stacking line is defined in a three-dimensional space, and a two-dimensional airfoil coordinate point set is generated in combination with the thickness distribution control rate; and determining a three-dimensional blade entity. According to the method, blade load distribution serves as a direct control target for suppressing flow noise and a vibration excitation source, and a blade thickness distribution reverse solution method with blade top load distribution as a target and geometric requirements and blade top leakage suppression as constraints is adopted. And establishing a set of analytic forward design system for directly generating the three-dimensional blade geometry from a preset performance target.

The present invention relates to a method for manufacturing a turbomachine blade, comprising the steps of: Generation (S1), by a neural network and from a reference value of a first parameter (P1) characteristic of the blade, of representative data for a set of candidate geometries for the blade, the data comprising, for each candidate geometry, a value of the first parameter (P1) satisfied by the candidate geometry and satisfying a proximity condition with the reference value; from the representative data generated by the neural network, determination (S2) of a blade geometry that best satisfies a value of a second parameter (P2) characteristic of the blade, the second parameter (P2) characteristic of the blade being different from the first parameter (P1) characteristic of the blade; manufacturing (S3) the turbomachine blade according to the determined blade geometry. Figure for the abstract: Fig. 1

The present disclosure relates to the technical field of aero-engine compressors, and particularly provides a method and device for evaluating aerodynamic stability of a compressor based on gas-solid two-phase flow, an electronic device and a storage medium. In view of the complex gas-solid two-phase flow inside an aero-compressor, the method for evaluating aerodynamic stability of the compressor based on the gas-solid two-phase flow comprises: constructing a control equation based on the gas-solid two-phase flow, wherein the control equation represents the influence of blade geometry on the flow field through a blade force source term and represents the influence of solid particles on the flow field through a particle force source term; constructing an eigenvalue equation according to the control equation; and evaluating the aerodynamic stability according to the eigenvalue equation. The technical solution of the embodiment of the present disclosure can solve the technical problem of huge calculation amount of aerodynamic stability evaluation under the gas-solid two-phase flow, and quickly and accurately evaluate the aerodynamic stability of the compressor.