155 results about "Compressor blade" patented technology

The invention belongs to the field of aero-engines and turbines, and provides a quick prediction method, system and device for a two-dimensional viscous compressible flow field of a gas compressor and a medium, and the method comprises the steps: obtaining CFD simulation data of a two-dimensional blade profile of the gas compressor under multiple working conditions, and carrying out the preprocessing to obtain a standardized training data set; on the basis of the data set, constructing a hybrid neural network model which takes a blade profile geometric parameter and an operation condition parameter as input and takes a two-dimensional viscous compressible flow field as output; combining mean square error, gradient loss and physical consistency loss based on RANS equation residual error to construct a total loss function, and performing end-to-end training on the model; during reasoning, to-be-predicted parameters are input into the trained model according to the same preprocessing mode, and then a flow field prediction result can be rapidly output. According to the method, the problems of long consumed time, low precision and poor consistency of traditional CFD simulation can be solved, millisecond-level flow field prediction is achieved, high precision and high physical consistency are achieved, and the method is suitable for efficient optimization design of the blade profile of the gas compressor.

The invention discloses an axial flow compressor blade anti-erosion optimization method facing a sand and dust environment, and belongs to the technical field of compressor blade anti-erosion design and optimization. According to the method, on the basis of a gas-solid two-phase flow erosion mechanism in a sand and dust environment, links such as working condition analysis, parameterization platform building, constrained multi-objective optimization method design, model building and optimization and index verification are integrated, and an integrated anti-erosion optimization process is formed. CFturbo, AutoGrid, CFX and Matlab are called in a unified mode through a Python script, automatic coupling of geometric modeling, grid division, gas-solid coupling simulation and an intelligent optimization algorithm of the blade is achieved, and collaborative optimization of aerodynamic performance and erosion resistance can be achieved under the constraint of erosion wear. Compared with a traditional method which depends on experience modification or single simulation, the method has the advantages that the optimization efficiency is remarkably improved while the simulation precision is guaranteed, the personal error risk is reduced, and the method can flexibly adapt to different compressor models and diversified operation conditions.

The invention provides an additive repair blade high-cycle fatigue limit prediction method based on random defect probability distribution, and relates to the field of aero-engine fan and compressor blade maintenance, and the method comprises the steps: obtaining the defect distribution of a real blade; establishing a prediction model based on a critical distance theory; according to the prediction model, critical distances of the multiple real blades are determined, and a first critical distance is generated; performing defect reconstruction based on the defect distribution of the real blade to generate a virtual blade; obtaining the first high-cycle fatigue strength of the virtual blade according to the defect distribution of the virtual blade, the first critical distance and the prediction model; respectively determining a plurality of defect parameters and a Pearson coefficient of the first high-cycle fatigue strength, and screening a first parameter significantly related to the first high-cycle fatigue strength; constructing a fatigue strength evaluation model; and inputting the first parameter into the trained fatigue strength evaluation model, and outputting the predicted high-cycle fatigue strength. According to the scheme, the high-cycle fatigue limit of the blade can be quickly predicted, and the prediction accuracy is improved.

The invention discloses an axial flow compressor blade profile optimization method based on main and auxiliary agent models and a related device, and relates to the field of compressor blade design, and the method comprises the steps: generating a plurality of design parameter groups based on to-be-optimized design parameters and a parameter adjustable range of a to-be-optimized axial flow compressor blade by using a data sampling algorithm, outputting a blade profile performance prediction value corresponding to each design parameter group by using the main and auxiliary combination agent model, screening a plurality of alternative design parameter groups based on a comparison result of each blade profile performance prediction value and an initial performance reference, and performing computational fluid dynamics simulation; and determining the alternative design parameter group corresponding to the optimal simulation performance parameter as a blade profile optimization parameter group. According to the method, the invalid risk of the design parameter group is reduced by using the adjustable range of the parameters, the performance prediction precision is improved based on the flow field characteristic parameters by using the main and auxiliary combination agent model, finally, the blade profile optimization parameter group is determined based on the simulation result of the screened alternative design parameter group, and the prediction precision is further improved.

The invention discloses a thermal flushing system design method for an aero-engine compressor blade, belongs to the field of aero-engine cleaning, and provides a thermal flushing system design method for removing salt deposited on the surface of the compressor blade for an aero-engine working under the marine environment condition. According to the design method, targeted cleaning of the cleaning fluid on the salt deposited on the surface of the blade of the gas compressor is innovatively provided, and the cleaning efficiency of desalting is improved. According to the method, a complete design process of the aero-engine thermal flushing system is formed.

The invention provides a gas compressor blade titanium fire friction test device and test method. The gas compressor blade titanium fire friction test device is used for testing the flame retardant property of a blade sample piece of a gas compressor, and is characterized by comprising a first rotating piece, a second rotating piece and a friction piece, wherein the friction piece is arranged on the first rotating piece; the blade sample piece is assembled on the second rotating piece, and the blade sample piece and the friction piece are arranged to be opposite to each other; the first rotating part can rotate in a first direction, and the second rotating part can rotate in a second direction opposite to the first direction; in the rotating process of the first rotating piece and the second rotating piece, the friction piece makes contact with the blade sample piece, and then the blade sample piece is rubbed. The test device can better simulate the working environment of the high-pressure compressor.

The present application relates to the field of gas turbine engine compressor test, and discloses a blade non-integer order vibration analysis method and system of multi-physical field dynamic signal, wherein the airflow excitation frequency and the circumferential mode order generated by unstable aerodynamic disturbance are obtained through dynamic pressure test, and the circumferential unstable flow mode characteristics causing blade non-integer order vibration are obtained; the blade vibration frequency and the pitch diameter are obtained through blade vibration dynamic test, and the test parameters of the two physical fields are comprehensively processed to realize the judgment of blade non-integer order vibration. The present application can effectively identify the occurrence of blade non-integer order vibration, can be used for the test of non-integer order vibration of aero-engine compressor blade and the mechanism analysis of related blade vibration failure fault, and has important significance for improving the overall safety and reliability of aero-engine compressor.

The invention discloses a blade surface identification and curved surface reconstruction method based on compressor blade deformation analysis, and belongs to the field of blade thermal deformation analysis. According to the method, the problems of thermal deformation and low accuracy in existing compressor blade reconstruction are solved. The method comprises the following steps: firstly, drawing a blade region as a structured grid, extracting a layer of grid on the surface of a blade, and defining a unit set and a node set; performing blade finite element conventional calculation, extracting corresponding data, and identifying a corner unit and a corner node as an initial unit; according to the topological relation of the units, the adjacent relation between the units is established, and a blade unit arrangement matrix is constructed according to the adjacent relation; determining inner and outer surfaces by using the arrangement matrix and the node relation in the unit; determining the surface of the blade according to the perimeter of a certain layer; and according to the node coordinates and the surface node matrix, expanding and writing into a point cloud format according to rows, and reconstructing a surface body. The method is used for blade thermal deformation reconstruction.

The application discloses a compressor blade phosphate aluminum coating precision grinding and polishing path generation method and system based on point cloud data processing, which comprises the following steps: obtaining point cloud data of the coating surface of the compressor blade by using a three-dimensional scanning device, filtering and denoising the point cloud data, and pre-processing the data, and aligning the pre-processed point cloud data with a robot machining reference coordinate system; based on the local curvature of the point cloud, the point cloud model is divided into multiple layers of contour slices by using a curvature adaptive slicing algorithm, wherein the slicing interval is negatively correlated with the local curvature radius; the contour points of each layer of slices are B-spline fitted to generate a continuous grinding and polishing path curve; the robot executes the grinding and polishing path curve, and integrates a real-time thickness detection device; when the thickness measurement value deviates from a preset threshold value, closed-loop feedback control is performed, and the grinding and polishing path and / or grinding and polishing parameters are dynamically compensated. The application has the advantages of high machining precision, high efficiency, high consistency and the like.

The invention discloses a gas compressor blade design method and device, electronic equipment and a storage medium, and the method comprises the following steps: S1, determining a boundary condition in the spanwise direction of a blade and a boundary condition in the flow direction of the blade, the boundary conditions in the flow direction of the blade comprise root section angle distribution, tip section angle distribution, root section thickness distribution and tip section thickness distribution; s2, closing the convective diffusion equation based on boundary conditions, giving a determined value of a control coefficient in the convective diffusion equation, and solving the convective diffusion equation to obtain angle distribution and thickness distribution of each section of the blade; compared with the prior art, the convection diffusion equation with clear physical significance is reasonably borrowed, the performance upper limit of the compressor blade is improved, meanwhile, the performance of the compressor blade is prevented from being reduced, practicability is high, and the method is suitable for wide application and popularization.

The invention belongs to the technical field of gas turbine gas compressor inspection and blade cleaning design, and particularly relates to a gas turbine gas compressor inspection and blade cleaning integrated structure which integrates a gas compressor inspection and blade cleaning structure on a gas compressor stator casing, so that holes in the gas compressor stator casing can be reduced, and the efficiency of gas turbine gas compressor inspection is improved. Compared with the prior art, large circumferential space is prevented from being occupied, interference between structures is reduced, the disassembly and assembly difficulty is reduced, the online cleaning function on the compressor blades is designed, the compressor blades can be kept clean and the pollution rate of the compressor blades can be reduced through daily online cleaning, the performance attenuation rate of the compressor can be further reduced, and the service life of the compressor is prolonged. The gas turbine can continuously work for a longer time, the maintenance cost is reduced, in addition, the cleaning nozzle is designed, the large spraying range and angle are achieved, the better atomization effect can be achieved, the cleaning efficiency and the cleaning effect can be improved, and deep cleaning of compressor blades can be achieved.

The application discloses a system identification method for mechanical damping of a compressor blade, and the method is as follows: a working condition is determined through steady CFD calculation of a multistage compressor, and is used as an initial field to carry out unsteady CFD calculation; blade modal force amplitude f and phase θ are calculated based on the results of the aerodynamic unsteady calculation and strength finite element analysis; aerodynamic damping under different pitch diameters is calculated by combining the influence coefficient method, and average aerodynamic damping coefficients of aerodynamic damping under all different pitch diameters are calculated; vibration response values of the blade under typical states in the test are measured, and are converted into blade generalized displacement q; the modal force amplitude f, the phase θ, the blade generalized displacement q and the average aerodynamic damping coefficients are brought into a formula to obtain mechanical damping of the blade. Advantageous effects: the real mechanical damping of the blade in a multistage environment is obtained, and based on the obtained mechanical damping, the prediction accuracy and reliability of forced response of the blade in subsequent blade type optimization and blade modification design of derivative types can be effectively improved.

The application discloses a compressor blade design method and device, electronic equipment and storage medium, including the following steps: S1: determining the boundary condition in the blade span direction and the boundary condition in the blade flow direction, wherein the boundary condition in the blade flow direction includes the root section angle distribution, the tip section angle distribution, the root section thickness distribution and the tip section thickness distribution; S2: closing the convection diffusion equation based on the boundary condition, giving the determined value of the control coefficient in the convection diffusion equation, solving the convection diffusion equation to obtain the angle distribution and thickness distribution of each section of the blade; compared with the prior art, the convection diffusion equation with clear physical meaning is reasonably borrowed, so that the performance upper limit of the compressor blade is improved, and the performance of the compressor blade is avoided from being reduced, the practicality is high, and the application is suitable for wide promotion and application.

The utility model relates to a special fixture for air conditioner compressor vane milling machine notch processing, including top head bottom plate one, top head bottom plate two, extruding piece, top head bottom plate one is connected with machine tool operation mesa through fixing piece, top head bottom plate two is detachably connected with top head bottom plate one through reciprocating friction cooperation, and extruding piece sets up on top head bottom plate two and is used for fixing work piece. The novel fixture provided by the utility model has simple structure, convenient operation, can effectively reduce the abrasion of machine tool mesa, prolongs machine tool service life, improves processing efficiency and precision simultaneously, and has higher practical value.

This invention relates to a dynamic balancing method for a flexible rotor system in an engine, belonging to the technical field of mechanical rotor dynamic balancing. It is applicable to the dynamic balancing of low-pressure rotors or power turbine rotors in turbofan engines when using flexible rotors. The turbofan engine includes a compressor blade disk and a hollow connecting long shaft. The dynamic balancing method includes: S1, inspection of the manufacturing process; S2, dynamic balancing tests on the connecting long shaft and blade disk respectively; S3, assembly and testing of the connecting long shaft and blade disk bearing the first identification on the test bench of the dynamic balancing device according to design requirements; S4, checking whether the dynamic indicators meet the standards; S5, full-range speed vibration testing of the connecting long shaft and blade disk as a whole on the platform of a rotor dynamics experimental apparatus; and S6, verification of the repeatability of assembly parameters and dynamic balancing, thus improving the efficiency of dynamic balancing.

The application provides a device for compressor blade length size three-coordinate detection, which comprises a plurality of compressor blade clamping units arranged in a matrix mode, and each compressor blade clamping unit mainly comprises positioning round pins, sliding U-shaped pressing blocks and U-shaped positioning limiting blocks. The three positioning round pins are arranged in parallel, one end of the positioning round pin comprises a spherical surface for supporting a side end surface of a compressor blade edge plate blade body; the sliding U-shaped pressing block comprises a first U-shaped notch for pressing a circumferential corner of the compressor blade edge plate; the U-shaped positioning limiting block comprises a second U-shaped notch for pressing the circumferential corner of the compressor blade edge plate; and the spacing between the sliding U-shaped pressing block and the U-shaped positioning limiting block is variable. The device can realize three-coordinate detection of multiple blades clamped at one time, improves the measurement efficiency and reduces the labor intensity of the inspectors.

The present application relates to a kind of robustness design optimization methods of data-driven compressor blade, using the way of middle camber line superposition thickness distribution to construct blade, using NURBS curve to parameterize middle camber line of blade.4-point 3-order data-driven non-embedded polynomial chaos method is used to quantify the uncertainty of sparse sampling data, and four blade configuration modes are obtained.Latin hypercube method is used to sample blade design space, and sampling set is used to train Gaussian process regression model under each blade configuration mode;GPR proxy model at each blade configuration mode is obtained respectively.After training, multi-objective optimization algorithm NSGA II is used to optimize search with the statistical mean and standard deviation of blade total pressure loss coefficient as target;Robust compressor blade with better performance and greatly reduced sensitivity to input uncertainty is obtained.

In order to solve the technical problems that an existing blade tenon spraying protection tool is not suitable for tenon bottom face spraying and a blade base body is prone to being scratched, the utility model provides the aero-engine compressor blade tenon cold spraying protection device. The blade tenon protection device comprises the sliding check block and the adjustable check block which are used for protecting the long side wall and the short side wall of the blade tenon respectively, and the protection requirement for a non-spraying area when the bottom face of the blade tenon is sprayed can be met. Moreover, a sealing rubber strip is arranged on the contact surface of the sliding stop block and the long side wall of the blade tenon, and a sealing rubber strip is arranged on the contact surface of the adjustable stop block and the short side wall of the blade tenon, so that the clamping and protection effects can be achieved through the sealing rubber strips, and the risk that the blade tenon is scratched due to clamping is avoided.

The utility model discloses a compressor blade angle zeroing device, and relates to the technical field of adjusting tools. The compressor blade angle zeroing device comprises a positioning plate and a blade angle zeroing device, a first positioning column and a second positioning column; the first positioning column and the second positioning column are both arranged on the positioning plate, and the first positioning column is matched with the first positioning hole; the second positioning column is matched with the second positioning hole; the positioning pipe is movably connected with the positioning plate; a threaded hole is formed in one end of the positioning pipe and matched with a threaded column arranged at the top end of the adjustable blade. Through the arrangement of the first positioning column, the second positioning column and the positioning pipe, the linkage ring in the gas compressor is located at the preset position, and the zero returning device can be installed in the gas compressor. In other words, if the zeroing device is installed on the air compressor, the linkage ring is certainly located at the preset position, namely, the adjustable blades corresponding to the linkage ring are located at the preset angle.

The invention provides a gas turbine compressor blade service life evaluation method based on high-cycle fatigue characteristics, and aims to solve the problem that fatigue tests under pure simulation and laboratory conditions cannot be effectively combined to meet the requirement of gas turbine blade type core component service life evaluation. According to the method, the strength condition and the vibration condition of the blade are determined through finite element analysis, then the stable attenuation law of the fatigue strength of the blade is obtained by actually measuring the vibration stress and the fatigue limit values of the blade in different use time periods, and the service life of the gas turbine compressor blade is comprehensively evaluated by taking the high-cycle vibration fatigue reserve as the standard. The method is not limited to a ship gas turbine compressor, is also suitable for effective evaluation of the service life of the moving blades of an industrial gas turbine and an aero-engine compressor, and can provide a fatigue reliability design direction for a new model.

The application relates to a supersonic rotor blade basic element airfoil design method, a rotor blade and an axial flow compressor, and belongs to the technical field of compressor blade design. Three sections of front, middle and rear are divided by adopting two shock wave surfaces in a rotor blade basic element channel, each section turning angle and relative Mach number are calculated, the rotor blade basic element airfoil is geometrically modeled based on the section turning angles, and the geometric modeling is verified and fine-tuned based on the relative Mach numbers, so that the problems of low efficiency and complex process of the supersonic blade basic element airfoil design method of the existing axial flow compressor are solved.

This disclosure relates to a method for designing compressor blades and the compressor blades themselves. The method includes: determining the suction surface profile of a basic blade profile with a blade height section, which includes: giving the leading edge metal angle β of the suction surface profile. in Trailing edge metallic angle β out And the axial projection length L of the suction surface profile, where β out <β in Establish a rectangular coordinate system OXY, with the X-axis parallel to the axial direction. The coordinate of the leading edge S of the suction surface profile is 0 on the X-axis, and the coordinate of the trailing edge C of the suction surface profile is L on the X-axis. Given the maximum turning point A of the suction surface profile, the coordinate of the maximum turning point A on the X-axis is greater than 90% and L is less than L. Given the metal angle β of the maximum turning point A. A ,β A <β out The suction surface profile is determined by making the metal angle of the suction surface profile gradually decrease from the leading edge point S to the maximum turning point A, and gradually increase from the maximum turning point A to the trailing edge point C. This disclosure is used to improve the aerodynamic performance of compressors and achieve precise control of the suction surface Mach number.

The invention relates to the technical field of engine testing, in particular to an aero-engine performance detection method, which comprises the following steps of: acquiring vibration signals mounted on an engine shell and blade tip timing data at the periphery of a gas compressor blade, and performing segmented processing on the continuous vibration signals according to an engine rotating speed period; and calculating time-frequency distribution of each signal segment to obtain a time-frequency energy matrix. According to the method, the vibration signals and the blade tip timing data are synchronously collected and segmented processing is carried out in combination with the rotating speed period, so that the frequency characteristics in different time periods can be completely extracted, and information missing caused by single time domain or frequency domain processing in a traditional method is avoided; furthermore, spectral density is calculated on a frequency domain and a cyclic frequency domain through a time frequency energy matrix, a two-dimensional spectrogram capable of reflecting energy distribution and cyclic characteristics at the same time is constructed, and full expression of characteristic information under multiple dimensions is guaranteed.

A turbomachine includes a compressor housing and a turbine housing. One of the compressor housing and the turbine housing includes a first seal component. The turbomachine also includes a unitary impeller rotatable about an axis, including a first face having a plurality of turbine blades facing in a first direction, and a second face having a plurality of compressor blades facing in a second direction. One of the first face and the second face includes a second seal component configured to cooperate with the first seal component to establish a fluid seal between the compressor housing and the turbine housing. The turbomachine further includes a shaft rotatable with the unitary impeller and an electric machine configured to convert at least one of rotational motion of the shaft to electrical energy and electrical energy to rotational motion of the shaft.

The application provides a heavy-duty gas turbine compressor aerodynamic strength vibration structure coupling design method and system. The method comprises the following steps: determining first flow-through parameters and first blade profile parameters of the compressor according to aerodynamic performance indexes; optimizing the first flow-through parameters and the first blade profile parameters of the compressor according to structural manufacturing and assembly indexes and the aerodynamic performance indexes, to obtain second flow-through parameters and second blade profile parameters of the compressor; optimizing blade root geometric parameters of the compressor according to strength vibration life indexes and the structural manufacturing and assembly indexes, and simultaneously designing and optimizing geometric parameters of a blade locking structure of the compressor, to obtain optimized blade root geometric parameters of the compressor and the geometric parameters of the blade locking structure; and designing the heavy-duty gas turbine compressor based on the second flow-through parameters and the second blade profile parameters. The technical scheme provided by the application improves the design efficiency, optimizes the performance balance, and reduces the design cost.

The invention relates to an aero-engine anti-icing and pneumatic cooperation method based on the bionic lotus leaf effect. The method comprises the following steps that micro-nano-scale super-hydrophobic surfaces are arranged in target areas of a fairing cone, an air inlet lip, the inner wall and the surface of the aero-engine and the surface of a fan / compressor blade of the aero-engine; the target area comprises a high-loss area of the front edge of the fan / compressor blade and the near tail edge of the suction surface; the front edge of the air inlet lip extends to an inner wall curvature abrupt change area and a fairing cone tip area; easy-to-freeze areas of the front edge of the fan / compressor blade and the tail edge of the suction surface; the micro-nano-scale super-hydrophobic surface has a micron / nano multi-scale structure, and is used for delaying icing and inhibiting ice layer accumulation. Compared with the prior art, the device has the function of realizing cross-scale cooperative regulation of pneumatic resistance reduction and icing and deicing; energy is saved, environment is protected, and low-energy-consumption anti-icing is And the pneumatic performance and the operation safety are improved.

This invention relates to a method for evaluating service damage of aero-engine compressor blades, solving the problem that current blade performance testing cannot obtain test values ​​'d' through direct sampling from the blade. The method comprises the following steps: S1: Design the dimensions of the blade simulator; S2: Select a base material consistent with the blade; S3: Calculate the maximum temperature and maximum stress of the blade under different service conditions, based on the blade's design status; S4: Statistically count the number of Category I low-cycle cycles, Category III flight takeoff cycles, and Category IV powered takeoff cycles for at least 10 engines, and calculate the conversion factors for each type of cycle using the temperature and stress calculation values ​​under different service conditions to obtain the equivalent number of cycles for different service times; S5: Apply low-cycle fatigue loading, fixing and clamping the simulator, and performing low-cycle cyclic loading according to the stated temperature, stress, and number of cycles; S6: Take samples from the low-cycle fatigue-treated specimen and perform tensile and high-cycle fatigue mechanical property tests. This invention significantly reduces the need for collecting blade service conditions and parameters, greatly reducing workload.

The invention discloses a gas compressor internal flow separation control method and device, and relates to the technical field of gas compressor flow control. According to the method, a working condition-excitation parameter mapping relation is established through database preprocessing, airflow parameters, flow state parameters and excitation state parameters are collected in real time through a multi-dimensional sensing system, the working condition is identified based on the airflow parameters, the flow separation degree is pre-judged through an LSTM algorithm, plasma excitation parameters are dynamically configured according to the pre-judgment result, and the flow separation degree is determined. And finally, parameters are optimized through closed-loop feedback, and the database is iteratively updated. The device comprises a compressor blade assembly, a modular plasma excitation system, a multi-dimensional sensing system, a self-adaptive control module and a working condition-excitation parameter database. According to the method, the problems that an existing scheme is limited in variable working condition adaptability and insufficient in control precision and response speed are solved, wide working condition coverage, pre-judgment and accurate control are achieved, energy consumption is reduced, and meanwhile pneumatic benefits and operation stability of the gas compressor are improved.

The application discloses a kind of inhibiting three-dimensional angle area separation compressor blade end area shaping method, comprising: machine casing end wall, wheel hub end wall and stator blade, the stator blade is arranged in ring between machine casing end wall and wheel hub end wall;N layers of machine casing side laminated grooves are arranged in the connecting side of stator blade suction surface and machine casing end wall;M layers of wheel hub side laminated grooves are arranged in the connecting side of stator blade suction surface and wheel hub end wall.The machine casing side laminated groove and the wheel hub side laminated groove effectively inhibit the end wall crossflow in high load compressor due to the strong adverse pressure gradient driving to climb up along the blade suction surface end area by imitating the feather laminated structure of bird wing, avoid the end wall crossflow to form low-energy fluid mass to block the blade passage under the action of flow direction strong pressure gradient due to separation.The three-dimensional angle area separation in high load compressor is effectively inhibited by regulating blade end area flow field, the total pressure loss is reduced, and the aerodynamic performance of compressor is improved.