33 results about "Compressor blade" patented technology

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 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.

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 application provides a novel vibration fatigue test method for a gas turbine compressor blade, which uses a material fatigue test statistical analysis method, uses a novel stress loading method on a blade test piece to excite blade resonance in a wide stress level coverage and a large number of stress level series, reasonably corrects stress levels according to a stress 'yield' principle, fills in the missing test stress sub-samples, reduces the dispersion of test data, and makes the stress sub-sample pairing more flexible and fast, and then obtains the median fatigue strength of the same batch of blades. The method can greatly save test cost, save test time, and avoid problems such as high test data variation coefficient or low confidence level caused by the structure itself characteristics or processing process error.

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.

The present disclosure relates to a robust design optimization method of compressor blade profile under multiple working conditions considering Mach number and a blade, the method comprising: establishing a design sample space by taking design parameters of the blade profile and incoming flow working condition parameters as variables; selecting multiple design sample points in the design sample space, each of which comprises the design parameters and the incoming flow working condition parameters; generating the blade profile corresponding to each design sample point according to the design parameters of the design sample point and a blade profile generation method; performing numerical calculation on the blade profile according to the incoming flow working condition parameters to obtain loss coefficients of the blade profile under different incoming flow attack angles; constructing a weight function and an objective function of the blade profile according to the loss coefficients and the incoming flow working condition parameters; establishing an approximate surrogate model according to the design parameters and the corresponding objective function values of each blade profile, and performing optimization calculation by using an optimization algorithm to obtain an optimal blade profile; the incoming flow working condition parameters comprise an incoming flow attack angle and a Mach number of the blade.

The application provides a design method of a compressor blade, and relates to the technical field of compressor blade profile design of gas turbines and aero-engine, and the method comprises the following steps: obtaining shape constraint parameters for designing the compressor blade, wherein the shape constraint parameters comprise a blade tip position, a blade root position, a blade tip deviation angle and a blade root deviation angle; determining a design shape of the blade according to a first position deviation degree of a first intersection point between a first reference line and a plane where the blade root position is located relative to the blade root position, and a second position deviation degree of a second intersection point between a second reference line and a plane where the blade tip position is located relative to the blade tip position; determining a plurality of control points matched with the design shape according to a third intersection point between the first reference line and the second reference line; and determining design parameters for designing the compressor blade according to the plurality of control points.

The application discloses an aero-engine compressor blade damage on-site repair method in the technical field of compressor blade damage repair, adopts any one or a combination of two of the repair methods of the repair method or the symmetric shearing method, removes the damage on the blade without disassembling the blade disc, reduces the influence of the damage on the compressor performance, simultaneously prevents the damage from further expanding, and the repaired blade disc can continue to be used, increases the service life of the blade disc, and guarantees the safety during the use of the blade disc.

The present application relates to a kind of compressor blade grid parameterization joint call analysis method.The present application is with MATLAB software as total control hub, by the flow control and data interaction ability of MATLAB, the grid parameterization deformation technology of HyperMesh software and the parametric finite element solution technology of ANSYS software are organically integrated together, construct a full-automatic closed loop analysis method and system from variable input to result output.The method realizes the quick response and batch calculation of several kinds of uncertain variables such as geometry, material, load, and is especially suitable for the fatigue life reliability analysis of compressor blade.

The application provides a compressor blade root mounting structure, a compressor and a blade root mounting method. The mounting structure comprises a wheel disc, a first groove body is arranged on the wheel disc and is distributed along the axial direction of the wheel disc, a second groove body is distributed along the circumferential direction of the wheel disc, a locking structure comprises a first locking block and a second locking block, the first locking block is provided with a first notch along the circumferential direction of the wheel disc, the second locking block is provided with a protrusion on the side close to the first locking block, in the first state, the protrusion fills the first notch, the other side of the second locking block is flush with the side wall of the first groove body, in the second state, the first end surface of the protrusion is in contact with the second end surface of the first locking block, the sum of the sizes of the first locking block and the second locking block in the circumferential direction of the wheel disc is greater than the size of the second groove body, the other side of the second locking block is in contact with the first blade, and the axial movement of the first blade is limited. The mounting structure provided by the application realizes the locking of the blade after the blade is mounted by means of the displacement of the second locking block, and the integrity of the flow channel is ensured.

The utility model provides an engine hot cleaning system, including cleaning liquid container, pressure supply device and nozzle, and cleaning liquid container communicates with the nozzle through the cleaning pipeline, and the nozzle is connected to the air compressor, and the cleaning liquid is sprayed into the air compressor passageway when the air compressor is operated in the slow operation state, and the cleaning pipeline is connected with the pressure supply device, the nozzle includes the connecting section, the neck -in section and the diffusion section who connects in proper order, the connecting section is connected with the cleaning pipeline, and the cross section of diffusion section is the fan shape, the utility model discloses the cleaning liquid is sprayed into the air compressor passageway when the air compressor is operated in the slow operation state, and the diffusion section of nozzle is made into the fan shape, makes the area of cleaning liquid spray as far as possible extensive contact with the air compressor blade, and after contacting, the continuous liquid phase is broken, atomized into small droplet by the shear disturbance effect of blade rotation, and as far as possible more air compressor blades are cleaned, so that more blades can be cleaned as far as possible under the condition of not disassembling the air compressor.

This application provides a method and storage medium for determining the installation angle of compressor blades, which can be applied to the field of aerodynamic design and performance optimization technology for compressors in aero-engines and gas turbines. The method includes: matching the input operating conditions of the compressor with a target correspondence to obtain the target installation angle of the compressor blades corresponding to the input operating conditions. The target correspondence is obtained by: acquiring multiple training sample pairs, each training sample pair including a set of operating conditions of the compressor and an installation angle of the compressor blades; using multiple training compressor performances corresponding to the multiple training sample pairs as labels for the multiple training sample pairs, and using the labels and the multiple training sample pairs to train an initial performance prediction model composed of multiple sub-models to determine the weights of the multiple sub-models, thereby obtaining a target performance prediction model; and determining the target correspondence based on the target performance prediction model.

This invention belongs to the field of aero-engine technology, specifically a method for assessing erosion damage to aero-engine compressor blades that integrates multiple operating parameters. The method includes: determining the fluctuation range of operating influencing parameters through on-site sampling and flight parameters; constructing a multi-parameter sample set using Latin hypercube sampling; deriving an erosion wear theoretical model; identifying key parameters related to target material characteristics and impact velocity by combining sand particle velocity measurement and titanium alloy erosion experiments; establishing a gas-solid two-phase flow finite element model; substituting the sample set into the model to solve for the wear rate concentration value; using a polynomial chaotic expansion method to obtain the relationship between operating influencing parameters and the wear rate concentration value; and deriving the statistical characteristics of the wear rate concentration value; finally, calculating the statistical characteristics of damage depth by combining erosion time and blade surface area. This invention achieves high-precision assessment under dynamic coupling of multiple parameters, solving the problems of single parameters and poor model adaptability in traditional methods, and significantly improving its engineering practical value.

The application discloses a slotted blade with variable attack angle jet flow automatic start-stop, which comprises a compressor blade, a bleed air port, a jet flow port, a jet flow slit, a movable valve, an anti-rotation elastic element, an elastic element groove, a stop step and a hinge; wherein the bleed air port is located on the blade basin side of the compressor blade, the jet flow port is located on the blade back side of the compressor blade, the jet flow slit is connected and penetrates the bleed air port and the jet flow port, the elastic element groove is arranged on the left side of the jet flow slit, the movable valve is installed in the elastic element groove through the anti-rotation elastic element and the hinge, and the stop step is arranged on the right side wall surface of the jet flow slit. According to the application, the jet flow can be automatically started and stopped according to different attack angles without introducing external air source or power supply, so that the external flow separation can be well inhibited under the large attack angle working condition, and the internal flow loss can be maintained to be low when there is no separation under the low attack angle.

The application discloses a flexible compressor blade with adjustable, which comprises a front blade frame and a rear blade frame arranged oppositely, a side plate fixedly connected to one side of the front blade frame and the rear blade frame, a supporting shaft arranged between the front blade frame and the rear blade frame, a flexible blade skin coated on the front blade frame and the rear blade frame, a supporting frame slidingly arranged on the supporting shaft, a small pulley near the front blade frame and a large pulley near the rear blade frame, the small pulley and the large pulley being fixedly connected through a supporting sleeve, the supporting sleeve being sleeved on the supporting shaft, a supporting belt being coated between the small pulley and the large pulley, and an adjusting and locking assembly being arranged between the supporting frame and the side plate. The application can realize continuous and smooth deformation, make the compressor fully utilize the lift effect and resistance effect of fluid to work, and make the engine reach the maximum aerodynamic efficiency, so as to meet the working conditions in a larger range and more severe conditions.

This invention belongs to the field of compressor flow field prediction technology, specifically involving a method and system for predicting the flow field of plasma-controlled compressor blades based on a physical information neural network. The method extracts the core input and output parameters of the flow field, integrates experimental and numerical simulation data and performs targeted preprocessing, constructs a pyramid-shaped multi-layer feedforward neural network, embeds the Navier-Stokes equations and boundary conditions, and trains the model using a composite loss function containing both data and physical losses. This invention balances high efficiency and physical reliability, achieving prediction accuracy comparable to traditional CFD with minimal computational resources. Even with scarce data, it can still output flow field results that conform to physical laws. It can quickly determine the optimal plasma excitation parameters, reduce the total pressure loss of the compressor, improve efficiency, and provide support for the optimization of aero-engine performance.

This utility model belongs to the field of material erosion testing technology, and particularly relates to an equivalent test device for compressor blade erosion. It includes a compressed air pump, a sand tank, an erosion pipe, an erosion chamber, and a sample holder. The sample holder is characterized by having a sample slot sealed on three sides, with a baffle on the open side of the slot. The sample holder has ear plates on both sides; one ear plate is movably connected to the side wall of the erosion chamber via a half-shaft, and the other ear plate is movably connected to the opposite side wall of the erosion chamber via a half-shaft. The outer end of either half-shaft is connected to the output shaft of a worm gear reducer, which is mounted on the outer side wall of the erosion chamber. The beneficial effects of this utility model are: by accurately controlling the quality of the abrasive material flowing over the surface of the flat material sample, it achieves precise simulation of the erosion effect, providing accurate data support for equivalent analysis, which is beneficial for guiding material optimization and ultimately extending the service life of compressor impeller products.

The application provides a proxy-assisted collaborative particle swarm optimization algorithm for expensive large-scale constrained optimization problems, and an urgent need for efficient processing of expensive large-scale constrained optimization problems in the field of intelligent manufacturing and the like. The existing methods often perform poorly when facing high dimensionality, high calculation cost and complex constraints at the same time. The application decomposes a complex optimization problem into a main-aid problem for collaborative solution by splitting large-scale decision variables, so as to overcome the curse of dimensionality. A hybrid proxy model management mechanism of the main-aid problem is designed, the model precision and the calculation resource allocation are dynamically optimized, and the calculation cost is reduced. A multi-population collaborative evolution optimization algorithm of the main-aid problem is given, and the solving efficiency is significantly improved by introducing an information migration and feedback mechanism. The method is applied to X benchmark test problems and an aero compressor blade aerodynamic optimization design problem, and is compared with x kinds of popular algorithms, and the effectiveness of the method is verified by experiments.

This invention relates to the field of compressor blade fouling, and particularly to a two-dimensional geometric uncertainty modeling method for compressor blade fouling. Based on the actual morphology of compressor blade fouling, the fouling is divided into a compact layer and a loose layer along the blade profile normal. Considering the uncertainty of fouling distribution on the compressor blade surface under actual operating conditions, a sparse compact layer fouling geometric uncertainty model is constructed by combining the KL expansion method with a sparse grid numerical integration method, and a randomized coarse loose layer fouling geometric uncertainty model is also established. The geometric uncertainty modeling method for compressor blade fouling provided by this invention is applicable to establishing geometric uncertainty models for compressor blades with arbitrary fouling distributions. The sparse blade fouling geometric model represents all geometric models within the entire distribution space, making it particularly suitable for providing CFD calculation geometric models for predicting the probability of aerodynamic performance degradation in fouled compressor blades.

FIELD: gas turbine engine.SUBSTANCE: method for strengthening compressor blades of a gas turbine engine.Shot-blast hardening of the blade profile is carried out using microbeads with a diameter of 0.2-0.4 mm. Subsequent vibration machining of the blade profile is performed in an abrasive environment. EFFECT: increase of the fatigue strength of the compressor blades while ensuring the required level of blade profile surface roughness.1 cl, 1 dwg

This invention relates to the field of turbomachinery technology and discloses a method for quantifying the uncertainty of aerodynamic performance due to geometric deviations caused by fouling on compressor blades. The method includes defining a sampling space for fouling variables and operating condition variables and obtaining an initial parameter sample set; generating a three-dimensional geometric model of the fouled blade based on the sample set; performing computational fluid dynamics simulation to extract flow rate, pressure ratio, and efficiency parameters to construct a sample database; training three independent fully connected neural network surrogate models to predict flow rate, pressure ratio, and efficiency using the database; sampling within the operating condition variable sampling space, inputting the flow rate surrogate model to obtain the predicted flow rate, and comparing the error with the target flow rate under a benchmark operating condition to obtain the optimal operating condition variable; inputting the parameter combination containing the optimal operating condition variable into the pressure ratio and efficiency surrogate models to output the predicted distribution, and completing the quantification of aerodynamic performance uncertainty through statistical processing. This invention reduces the computational cost of large-scale evaluations and improves the efficiency of quantifying aerodynamic performance degradation under complex fouling conditions.

The application provides an electric air compressor (1), which avoids the reduction of the efficiency of the compressor by eliminating the gap between the front end of the blade (22) of a vane diffuser (20) and a wall surface (1131) of a compressor shell (11). The electric air compressor (1) comprises a compressor (10), a vane diffuser (20) and an electric motor (30) which are rigidly connected in sequence, and the compressor shell (11) is axially butted with the vane diffuser (20) in a manner of abutting the front end of the blade (22) of the vane diffuser (20).

This invention discloses a biomimetic wavy trailing edge design method for compressors, belonging to the field of compressor design. The method includes: acquiring an original compressor blade model and constructing its arc function; adjusting the axial chord length distribution of the basic cross-sections at different spanwise positions using a waveform control function to form a trailing edge geometry that varies wavyly along the spanwise direction; calculating the suction side thickness distribution, pressure side thickness distribution, and the angle between the normal to each side profile and the axial direction for each basic cross-section of the prototype blade; performing coordinate transformation on the arc function, and superimposing the suction side thickness distribution, pressure side thickness distribution, and the angle between the normal to each side profile and the axial direction onto the transformed arc function while maintaining these parameters unchanged, generating the biomimetic suction and pressure surface coordinates, thereby forming a flow guide channel on the suction side of the blade that is continuous with the wavy trailing edge geometry; and generating a biomimetic wavy trailing edge blade based on the adjusted basic cross-section geometry. The structure of this invention can effectively suppress corner separation and reduce flow losses.

This utility model belongs to the field of centrifugal compressor manufacturing technology, and particularly relates to an anti-erosion and wear structure for compressor blades. Its key feature is that it includes ceramic fiber cloth, which is integrally bonded to the blade body surface using a silicone rubber adhesive. The ceramic fiber cloth is woven with a herringbone pattern or has a herringbone pattern hot-pressed onto the surface of the ceramic fiber felt using a mold, forming a symmetrical V-shaped groove surface similar to shark skin. Compared with existing technologies, the beneficial effects of this utility model are: by bonding a layer of ceramic fiber cloth with a shark skin-like texture to the surface of the aluminum alloy blade, the unique friction-reducing effect of the shark skin texture is utilized to reduce the erosion effect of the fluid on the impeller surface; the ceramic fiber cloth material itself has wear resistance, high temperature resistance, and corrosion resistance, which can improve the defects of the aluminum alloy material itself, extend the service life of the impeller, and ensure the long-term safe and stable operation of the centrifugal compressor.

The present invention relates to a compressor blade for a gas turbine engine. In particular, the blade of the compressor is modified according to predetermined requirements of both aerodynamic stability and fuel economy in a plurality of planes.

This invention discloses a high-dimensional optimization method for compressor blades based on multi-model dimensionality reduction and sensitivity guidance, belonging to the field of compressor blade design. The method includes dimensionality reduction of the compressor blade design variables, generating samples based on the key variables of the dimensionality reduction, training a surrogate model, and calculating the total effect Sobol exponent of the isentropic efficiency and pressure ratio for each key variable based on the output of the surrogate model. Based on the total effect Sobol exponent, a population and individual crossover mutations are generated. Then, it is determined whether to update the surrogate model. If not updated, or if updated, a non-dominated sorting is performed based on the objective function values ​​of individuals in the new population to obtain a frontier set. A new generation population P is then constructed from the frontier set. g Determine if g equals T; if so, then in P... g Candidate solutions are selected from the non-dominated solution set and input into the CFD solver. Samples that meet the design requirements in terms of isentropic efficiency and pressure ratio calculated by the CFD solver are selected as the final sample set; otherwise, the process returns to the crossover and mutation step.

An aircraft powerplant assembly includes a compressor rotor, a compressor vane structure, a diffuser structure, a combustor and an air system. The compressor rotor includes a rotor stage and a rotor hub. The rotor hub includes a hub wall connected to a rotor base. An outer surface of the hub wall forms an inner peripheral boundary of a first section of an air passage. The compressor vane structure is radially outboard of the hub wall. The diffuser structure includes a diffuser nozzle and a diffuser plenum. The combustor is disposed in the diffuser plenum. The air system includes an air circuit and a heat exchanger. The air system is configured to bleed compressed air from the diffuser plenum through a circuit inlet, cool the compressed air with the heat exchanger to provide cooled air, and direct the cooled air into the air passage through a circuit outlet.

The application provides a preparation method of a wear-resistant hard coating based on ion implantation and ion nitriding on a TC4 titanium alloy surface, and comprises the following steps: vacuumizing a TC4 titanium alloy sample and then performing ion implantation, inputting mixed gas of hydrogen and argon, and starting glow discharge; changing the mixed gas of nitrogen, hydrogen and argon, and starting ion nitriding; vacuumizing again, heating, adjusting the nitrogen flow, keeping vacuum state, opening a Cr target arc power source to deposit a CrN layer to form a primer layer; adjusting the flow of the mixed gas of nitrogen and argon, keeping vacuum state, depositing a Cr layer, then adjusting the nitrogen flow, and depositing a CrN layer; alternately depositing the Cr layer and the CrN layer; taking out the workpiece after the workpiece is cooled to room temperature in the furnace; and the application provides a wear-resistant hard coating based on ion implantation and ion nitriding on a TC4 titanium alloy surface, which is suitable for surface strengthening of key components such as compressor blades of an aero-engine, artificial joint implants, marine engineering fasteners and valve of nuclear power equipment.