122 results about "Tip clearance" patented technology

The invention belongs to the technical field of design of a cold-state blade tip gap between an aero-engine rotor blade and a cartridge receiver, and particularly relates to a method for determining the cold-state blade tip gap between the aero-engine rotor blade and the cartridge receiver. According to the method, the steady-state point is calculated, including the overall flexural deflection of the rotor, the local variation caused by rotor and stator machining runout, the rotor bearing clearance and the like, and transition state calculation is added while steady-state point calculation is performed, so that the actual use scene of the aero-engine can be better considered, and the long-term use requirement is met.

The invention discloses a calculation method of a fan blade tip gap accelerated flow analytical model, and belongs to the technical field of fan speed field calculation. By analyzing the aerodynamic characteristics of a fan wake flow gap acceleration area, the correlation between the power increase of the side-by-side fan and the fan wake flow gap acceleration area is disclosed, and an analytical model is constructed according to the self-similarity of the gap acceleration area and an eddy current equation. According to the model, local accelerated flow caused by rotation of adjacent rotor blade tips under the same plane is considered, by adopting the method, the used model has high precision in the aspect of predicting the flow field characteristics of the blade tip wake flow gap acceleration area, the prediction effect is close to that of a high-fidelity numerical simulation method, but the calculation cost is far lower than that of traditional CFD simulation; the method is well applicable to processing and simulating wind power plant structures which are closely arranged along the y-z same plane, such as a double-head floating fan and a fan wall, and shows good engineering application value.

The invention discloses a cross-scale multi-physics-field coupling aero-engine warming-up process analysis method. The method comprises the steps that firstly, an engine air system one-dimensional network model, a fuel and lubricating oil system one-dimensional network model, a wheel disc and cartridge receiver two-dimensional axial symmetry model, a blade parameterization model and a main runner quasi-one-dimensional model are established; then calculating an air system boundary and a fuel and lubricating oil system boundary by using a main runner cavity method, then calculating the fuel and lubricating oil system boundary and the air system boundary by using a one-dimensional transient fluid network method to obtain a heat exchange boundary of a wheel disc and a cartridge receiver, and considering a thermal load and a centrifugal load in an unsteady state heat conduction process; calculating the radial displacement of the wheel disc and the casing by using a two-dimensional axisymmetric finite element method, and calculating the temperature and the radial displacement of the blades by using an engineering algorithm; and finally, integrating the radial displacement of the wheel disc, the blades and the casing and the initial blade tip clearance value to obtain the blade tip clearance value of each stage in the transient process, and feeding back the blade tip clearance value to the mainstream model again to realize iterative calculation. The analysis precision of each physical field is improved, and the calculation time consumption is reduced.

The invention discloses an aero-engine blade data storage method, device and equipment and a medium, and relates to the technical field of data processing. The method comprises the steps that signals of all blades of the aero-engine are collected through a plurality of sensors, and blade vibration signals and blade tip gap signals corresponding to the sensors are obtained; generating health characteristic parameters corresponding to the sensors according to the blade vibration signals and the blade tip gap signals corresponding to the sensors, and generating signal auxiliary information corresponding to the sensors according to the health characteristic parameters; determining a time cache granularity corresponding to each sensor according to each piece of signal subsidiary information, and allocating a continuous address cache region corresponding to each sensor; and storing a blade vibration signal, a blade tip gap signal, a health characteristic parameter and signal affiliated information corresponding to each sensor according to each time cache granularity and each continuous address cache region. In this way, efficient and low-redundancy storage can be achieved, and high real-time performance and flexible expansion are achieved.

The invention relates to the technical field of fluid machine design and artificial intelligence crossing, in particular to an optimization design method and system oriented to turbomachinery, a storage medium and a computer program. According to the method, a standardized data bus of a parameter layer, a configuration layer, a geometric layer and a result layer penetrates through an average line, parameterized geometry and multi-objective optimization are carried out, thermophysical property calibers are unified, and sample-level archiving and playback are achieved. Through LHS / Sobol sampling, Kriging proxy, NSGA-II and other algorithms, isentropic efficiency, pressure ratio and structure / noise indexes are collaboratively optimized under the condition that constraints of mass flow, minimum thickness, tip clearance and the like are met; the system is coupled with CFD / FEM through an adapter, and supports GUI and batch processing. Compared with the prior art, the design efficiency, the cross-module consistency and the result traceability are improved, and a manufacturable Pareto optimal blade scheme is obtained.

The invention relates to the technical field of aero-engine state monitoring, and discloses an aero-engine blade tip gap measuring probe and system based on an eddy current method.The aero-engine blade tip gap measuring probe adopts a coplanar Archimedes spiral double-coil array, and in the measuring process, the blade tip gap can be measured by applying anti-phase excitation; the common-mode noise suppression in the blade tip clearance measurement process is realized by combining a differential amplification circuit; signal extraction and temperature drift compensation are realized through a phase-locked amplification technology and an impedance-resistance orthogonal decoupling architecture, and finally micron-order gap analysis is realized. The aero-engine blade tip clearance measurement system provided by the invention can still keep excellent reliability under extreme working conditions of high temperature, high rotating speed and the like of a turbine, and meanwhile, the measurement precision of the blade tip clearance can be remarkably improved.

The invention provides a high-pressure turbine outer ring connecting and positioning structure capable of being used for adjusting blade tip gaps, and belongs to the field of aero-engines. The outer ring fixing ring is fixed on the turbine case; the high-pressure turbine outer ring is mounted on the outer ring fixing ring through a connecting pin; the connecting pins are provided with pin heads with circumferential positioning surfaces, large-diameter matching sections and small-diameter matching sections, the circumferential positioning surfaces are matched with planes or arc surfaces on the outer ring fixing ring to realize circumferential positioning, the large-diameter matching sections are matched with mounting holes in the outer ring fixing ring, and the small-diameter matching sections are matched with pin holes in the outer ring of the high-pressure turbine; the connecting pin comprises a non-eccentric structure and an eccentric structure with a small-diameter matching section facing downwards or upwards at different degrees; when the outer ring of the high-pressure turbine is assembled on the outer ring fixing ring, the radial height of the outer ring of the high-pressure turbine is adjusted by selecting and matching the connecting pins with different eccentricity as required, so that the blade tip clearance between the blade and the outer ring of the high-pressure turbine is adjusted.

The invention discloses a turbine blade tip clearance passive matching method based on engine requirements, and belongs to the technical field of aero-engines, and the turbine blade tip clearance passive matching method comprises the steps that the minimum allowable working clearance is determined; determining a typical tip clearance analysis load spectrum; performing thermal analysis and deformation analysis on the rotor and the stator; according to the deformation results of the rotor and the stator, a blade tip gap change rule is analyzed, and a blade tip gap change rule is obtained; according to different engine requirements, blade tip clearance matching is optimized according to the material, the structure and the cooling form of the cartridge receiver; determining an initial cold-state gap according to the narrow-point gap and the minimum allowable working gap; and carrying out iterative calculation, and determining a final blade tip clearance matching scheme. Turbine rotor deformation and turbine stator deformation can be matched so as to ensure that the turbine blade tip clearance is as small as possible in the focused operation state, that is, the turbine blade tip clearance is reduced as far as possible, the turbine efficiency is improved, and the oil consumption rate is reduced on the premise of ensuring the safety of an engine.

The invention discloses a rotor blade tip gap design method based on cold and hot state transformation and an aero-engine. The design method comprises the steps that the design working condition of cold and hot state transformation of a fan rotor is determined; performing pneumatic calculation on fan components; designing the structure of the hot-state blade S1; reversely solving the hot-state blade S1 to obtain a cold-state blade S2; designing a cold-state blade structure; performing forward solution on the cold-state blade to obtain a hot-state blade S3; forward solving is carried out on the fan rotor cold-state blade S2 to obtain a corresponding hot-state blade S3; performing error analysis on the thermal-state blade S3 and the thermal-state blade S1; blade tip models of the cold-state blade S2 and the hot-state blade S1 are compared and analyzed, and the front edge radial deformation delta q and the tail edge radial deformation delta h of the fan rotor are obtained; and according to the designed thermal state gap value delta r, the front edge gap Sq and the tail edge gap Sh of the blade tip of the fan rotor blade are obtained.

The invention discloses a synchronous decoupling measurement method for vibration displacement and blade top clearance of a shrouded blade, and relates to the field of energy power engineering, and the method comprises the following steps: 1, calibrating the axial vibration displacement and blade top clearance of the shrouded blade; step 2, installing an eddy current sensor; 3, the rotor is made to operate normally, the vibration state of the shrouded blade is measured in the rotating process of the shrouded blade, and a measurement signal is obtained; step 4, carrying out preprocessing on the measurement signal; 5, the sensing area variation of the sensor and the measurement distance variation from the sensor to the shroud are measured; and 6, calculating the vibration displacement and the blade top clearance of the shrouded blade. According to the method, the defects of the prior art can be overcome, synchronous and decoupling measurement on the vibration displacement and the blade top clearance of the shrouded blade can be realized, the measurement accuracy and credibility can be effectively improved, and the measurement requirement can be effectively met.

A turbine blade and related method are disclosed. The blade includes an airfoil body having a radial outermost surface relative to a turbine rotor, a bond coat over the radial outermost surface, and an outermost abrasive layer over the bond coat. The outermost abrasive layer may include a graded oxide ceramic layer or a graded metal ceramic layer. In any event, the outermost abrasive layer is vacuum heat treated to harden the layer, resulting in increased rub and improved clearance between the blade tip, creating a smaller tip gap and lower leakage compared to conventional systems. The heat-treated abrasive layer is also resistant to oxidation at higher temperatures, e.g., greater than 1090° C. (˜2000°F.), caused by the increased use of higher temperature fuels, such as hydrogen. The vacuum heat treatment also results in thermal cycle testing improvements.

The application discloses a kind of magnetic fluid sealing rim impeller structures for inhibiting blade tip clearance leakage, including pump shell, be provided with blade in pump shell, a circle of blade cover plate is equipped in the periphery of blade, blade cover plate and pump shell have sealing gap in radial direction, one or more magnetic fluid sealing rings are arranged in sealing gap along the axial direction;Magnetic force structure is arranged in the outer wall of pump shell in annular, magnetic concentration structure is arranged in the inner wall of pump shell in annular, closed magnetic circuit is formed between magnetic concentration structure and magnetic force structure, so that gradient magnetic field is formed at sealing gap;Magnetic fluid is adsorbed and positioned in the corresponding area of sealing gap under the action of gradient magnetic field, and forms a magnetic fluid sealing ring which is consistent with the shape of sealing gap, continuous and stable.The application utilizes the self-positioning characteristics of magnetic fluid under magnetic field, forms dynamic, self-adapting sealing interface between blade cover plate and pump shell, realizes zero leakage operation, improves the hydraulic efficiency and stability of axial flow pump unit.

A fan assembly has a rotatable propeller having a plurality of fan blades, each having an outer tip. A cylindrical fan shroud surrounds the propeller and has an inner surface facing the propeller. The inner surface comprises a pocket surface longitudinally aligned with the propeller and a constricted exhaust surface longitudinally downstream of the pocket surface and radially offset from the pocket surface by a pocket depth. Outer tips of the fan blades may define a tip clearance gap between the propeller and the inner surface. The outer tips may extend into an annular pocket. An effective tip clearance gap, defined by a distance between a diameter of the constricted exhaust surface and the outer tips of the fan blades, is smaller than a tip clearance gap between the outer tips and the pocket surface. Related fan assemblies for agricultural building are also disclosed.

A lead cooling assembly includes a cooling jacket including a tubular section having an exterior portion and an interior portion, where the interior portion is proportioned to surround a probe lead and provide space for a flow of a gaseous coolant. The lead cooling assembly also includes an intake configured to admit the gaseous coolant into the interior portion of the cooling jacket. The lead cooling assembly further includes a collar configured to fit over the exterior portion of the cooling jacket, where the collar includes one or more exit holes configured to pass the gaseous coolant out of the interior portion of the cooling jacket.

An aircraft engine includes a gas path annularly extending around a central axis in serial flow communication across a compressor and turbine. The turbine has a casing, a shroud secured to the casing, and blades radially inward of the shroud, with a clearance defined between blade tips and the shroud. An active tip clearance control device for directing a flow of cooling air to the casing includes a first cooling air source at a first temperature, a second cooling air source at a second temperature than the first temperature, and a valve having a first inlet fluidly connected to the first air source, a second inlet fluidly connected to the second air source, and an outlet selectively connectible to the first and second inlets. The outlet is fluidly connected to the casing to selectively direct the flow of air from the first or second air source to the casing.

Proposed are a blade tip clearance sensor and a gas turbine including the same. The blade tip clearance sensor includes a bushing portion which is fixed to a ring segment mounted on a turbine casing that accommodates a turbine blade therein and which has an internal space, a washer portion mounted on a lower portion of the internal space and provided with a first hollow portion, a flange portion provided on an upper surface of the washer portion and provided with a second hollow portion, a sensor probe which is inserted into the first and second hollow portions and which has a sensor cable, a spacer provided on an upper surface of the flange portion and configured to accommodate the sensor probe therein, and a locking screw which is provided on an upper surface of the spacer and which has a thread engaged with a thread of the bushing portion.

The invention belongs to the technical field of aero-engines, and particularly relates to a fan rotor blade structure and a design method thereof. The blade back side damping rib is arranged on the back side of the blade body of the basic blade; the blade basin side vibration reduction ribs are arranged on the blade body basin side of the basic blade; and the outer contour lines of the blade back side vibration reduction ribs and the blade basin side vibration reduction ribs are formed by outwards offsetting the blade body basic blade profile of the cross section of the basic blade at the same height position. According to the application, the aerodynamic performance loss caused by blade tip chamfers can be eliminated, and the structural safety risk caused by too small blade tip gaps is reduced; the problem of non-whole-order vibration of the blade tip of the fan rotor blade can be effectively solved, and the modal inherent frequency of each order of the rotor blade is improved by adjusting the structural parameters of the vibration reduction ribs on the two sides of the blade body.

The invention relates to the technical field of hydraulic machinery, in particular to a controllable radial thickening T-shaped blade top design method, an impeller and a blade. The method comprises the following steps: establishing a mapping blade of an original blade under a blade affine coordinate system; determining a closed two-dimensional geometric boundary formed by a plurality of curves in a relative circumferential-relative chordwise cylindrical surface of the mapping blade, and adjusting the closed two-dimensional geometric boundary based on a preset thickening parameter to obtain a thickened closed two-dimensional geometric boundary; and based on the thickened closed two-dimensional geometric boundary, determining a T-shaped blade top mapping blade under the blade affine coordinate system, and reflecting the T-shaped blade top mapping blade under the blade affine coordinate system to the original impeller cylindrical coordinate system to obtain a T-shaped blade top blade under the original impeller cylindrical coordinate system. By the adoption of the scheme, accurate control over the streamline normal direction can be achieved, the blade top clearance leakage vortex is restrained, and meanwhile the method has the functions of structure parameterization, control comprehensiveness and optimization flexibility.

The present invention belongs to the technical field of aircraft secondary power systems and discloses a rotor tip clearance measurement method and device. The method comprises the following steps: first, measuring the distance a between the outer ring surface of the inlet outer ring support and the end face of the inlet outer ring, as well as the height of the inlet outer ring end face on a plane, to obtain the distance c from the bottom surface of the inlet outer ring to the clearance design point; then, directly measuring the distance between the lower end face of the product's intake bearing end cover and the surface of the intake bearing end cover for assembling the L-shaped sealing gasket, and measuring the distance between the bottom of the inlet outer ring and the lower end face of the intake bearing end cover. The distance A between the clearance design point on the outer ring of the inlet duct and the surface of the intake bearing end cover for assembling the L-shaped gasket is calculated; the axial clearance f of the product bearing under preload is then measured; finally, by measuring the thickness h of the L-shaped gasket and the distance g between the clearance design point on the centrifugal impeller and the bottom of the bearing inner ring, the distance B between the clearance design point on the impeller and the end face of the L-shaped gasket is obtained, B = g + the theoretical distance between the end faces of the inner and outer rings of the bearing ‑f / 2 + h; the tip clearance C at the clearance design point on the centrifugal impeller is calculated as B‑A. Compared with existing measurement methods, the method of the present invention measures the tip clearance during product assembly, eliminating the need for repeated disassembly and assembly of the entire product and the need to transfer the product back and forth between measuring equipment. This improves product assembly efficiency and is more suitable for batch operations.

An aircraft engine includes an annular gas path extending across a compressor and a turbine with a casing, a shroud and blades, a clearance defined between blade tips and the shroud. An active tip clearance control device directing a cooling airflow to the casing includes a first air source at a first temperature, a second air source at a second temperature greater than the first, a valve having first and second inlets connected to the sources and an outlet selectively connectible with the inlets. The outlet is fluidly connected to the casing to selectively direct the flow of cooling air from the first or second source. A sensor measures the clearance. A controller connected to the sensor receives a signal therefrom of a measured clearance, and generates commands causing the valve to open to the first or second source when the clearance is above or below a predetermined value.

Disclosed herein is a gas turbine including a turbine casing and a turbine rotor arranged for rotation in the turbine casing. The gas turbine further includes a stator shroud concentric with the rotor and positioned radially outward from the blade tips. The stator shroud comprises a plurality of shroud segments. A clearance control assembly, is provided, which includes a plurality of actuators that are functionally coupled to the shroud segments. Each actuator comprises: a bimetallic element, adapted to displace an actuator rod in a radial direction from the first radial position to a second radial position, through deformation of the bimetallic element caused by a temperature variation generated by a heater. The actuator rods are connected to the shroud segments and control a radial displacement thereof.

Systems and methods for a gas turbine engine include a rotor having a rotor blade that has a base, a tip, and an outer surface defined by a leading edge, a trailing edge, a suction side wall, and a pressure side wall, all extending between the base and the tip. Geometry of the tip of the rotor blade has a contour that is directed first toward the suction side wall at the leading edge, and then toward one of the pressure side wall or the suction side wall at the trailing edge. The geometry has a maintained or contracted shape that is shifted toward the suction side wall. The geometry is configured to reduce tip leakage interaction of the rotor to delay onset of tip vortex generation and desensitize the rotor to tip clearance.

The application discloses an aero-engine tip clearance testing system and a testing method. The system comprises a base, a driving motor, a shaft coupling, first to third rotating shafts, first to second supporting members, first to second bearings, first to second engine simulation modules, first to second double optical fiber sensors, first to second rotating speed sensors, a collector, a three-degree-of-freedom displacement platform, a fixing rod and a control module. The application has a two-stage rotor impeller and a casing, can measure the tip clearance of multiple sections, and has good mobility and displayability. The application can meet the tip clearance measurement and calibration of multiple sections, multiple impellers and multiple measuring points under different rotating speeds and different tip clearances, and the continuous adjustment precision of the tip clearance can reach 0.02 mm. Meanwhile, the application has small volume, good mobility and displayability.

The application aims to provide a compressor aerodynamic performance analysis method, an analysis system and a medium. The compressor aerodynamic performance analysis method comprises the following steps: obtaining a design value of the overall efficiency of the compressor; obtaining an over-limit value and a design value of the tip clearance of the rotor blade of any specified stage of the compressor; obtaining an efficiency difference value when the tip clearance is the over-limit value and the design value; calculating a third efficiency value when the tip clearance is the design value, so that the actual efficiency is equal to the difference between the third efficiency value and the efficiency difference value, and recording the outlet temperature calculation amount; obtaining the compressor outlet temperature calculation amount according to the outlet temperature calculation amount; and obtaining the actual overall efficiency of the compressor when the tip clearance of the rotor blade of the specified stage is the over-limit value according to the compressor outlet temperature calculation amount. The compressor aerodynamic performance analysis method can improve the efficiency and accuracy of analyzing the influence of the tip clearance change of the multistage axial flow compressor on the performance.

The application discloses a kind of centrifugal compressor impeller total play clearance pre-adjustment method, this pre-adjustment method can carry out total play clearance prediction when not installing big cover, only need to utilize the pressing device made of casing connecting bolt hole, instead of big cover and carry out pressing, and the gasket made for stress uniformity, total play clearance prediction can be completed, and if total play clearance does not meet requirements after big cover is reloaded, it is effectively avoided.The application overcomes the problem that the workload is obviously increased due to the disassembly of the big cover, improves the work efficiency, reduces the labor intensity, shortens the maintenance period, and ensures the maintenance quality.

The application discloses a roughened casing wall configuration for suppressing tip clearance cavitation by using the wall attachment effect. The configuration is uniformly arranged on the inner wall of the casing in the circumferential direction and covers the blade tip region in the axial direction. After the leakage jet is formed, the leakage jet will enter the configuration by using the wall attachment effect and form an induced vortex in the configuration to suppress the development of the leakage jet. The configuration is designed by using the flow characteristics of the leakage jet in the gap and the formation mechanism of the tip clearance vortex system, so that the turbulent flow is triggered early in the gap, the viscous dissipation of the leakage jet is increased, and the stable suppression effect is achieved at the source of the vortex system and cavitation.

The present application is suitable for the field of gas turbine technology, and provides a gas turbine blade tip dynamic gap prediction method and device. In the embodiment, first, a temperature field calculation method of the blade, turbine disc and turbine holding ring, and an initial blade tip gap calculation method are established, then the actual overall thermal parameters, metal wall surface temperature, rotating speed, rotor axis position and blade tip gap value of the gas turbine are obtained through sensors to correct the above calculation methods, and the dynamic change of the blade tip gap is determined through the corrected model and transient parameters under high rotating speed and high load. This solves the problem that the existing blade tip gap control technology based on the control strategy of static blade tip gap margin is prone to accidents under high rotating speed and high load transient conditions. And the gap value under the transient condition is predicted, which can effectively reduce the risk of blade tip rubbing, reduce the blade structure damage caused by rubbing, improve the operation safety of the gas turbine, and avoid the loss caused by unplanned shutdown.

The application discloses a semi-open type radial flow and mixed flow impeller tip clearance checking design method and device, relates to the technical field of impellers, and comprises the following steps: obtaining basic data and setting parameters of an impeller tip, wherein the basic data comprises tip deformation data, cold-state tip profile data and meridian cover line data; constructing a curve coordinate mapping relationship of cold-state-hot-state tip profile based on the tip deformation data and the cold-state tip profile data, and constructing a mapping relationship of the curve coordinate of the cold-state tip profile and meridian coordinates based on the cold-state tip profile data and the meridian cover line data; calculating the cold-state tip clearance distribution and the hot-state tip clearance distribution of the pressure side and the suction side tip line of a target blade changing with meridian coordinates according to the curve coordinate mapping relationship of the cold-state-hot-state tip profile and the mapping relationship of the curve coordinate of the cold-state tip profile and meridian coordinates; and calculating the cold-state clearance distribution of the target blade based on the cold-state tip clearance distribution, the hot-state tip clearance distribution and the setting parameters.

The invention belongs to the technical field of rotor dynamics, and discloses a blade tip clearance rapid prediction method in a maneuvering flight state. Additional load excitation and additional parametric excitation caused by maneuvering flight are derived based on the beam unit and the concentrated mass. A finite element model of the birotor-bladed disc-cartridge receiver coupling system is established based on a beam-shell hybrid model, additional load excitation and additional parameter excitation caused by maneuvering flight are fused, the finite element model of the birotor-bladed disc-cartridge receiver coupling system in a maneuvering flight state is finally established, and gaps of blade tips are calculated through a numerical integration method. A blade tip gap prediction model is constructed based on an artificial neural network, model training is completed based on a numerical calculation result serving as a training set, and finally rapid gap prediction under different working conditions is achieved. According to the method, the influence of centrifugal deformation of the blade, vibration displacement of the rotor, maneuvering flight deformation and other factors on the blade tip clearance is effectively considered, the morphology of the complex blade is comprehensively considered, and the model generalization ability is high.

The present application relates to the technical field of centripetal turbine optimization design, and particularly discloses a centripetal turbine optimization design method based on an advanced optimization algorithm, which comprises the following steps: S1: a one-dimensional aerodynamic thermodynamic mean diameter design model is constructed, and the model is integrated with loss models including flow passage loss, blade tip clearance loss, tail edge loss and outlet kinetic energy loss; S2: the one-dimensional mean diameter design model of step S1 is coupled with a particle swarm optimization algorithm with the maximum total static efficiency of the centripetal turbine as the target, and global optimization is performed on dynamic parameters, thermodynamic parameters and geometric parameters affecting the performance of the turbine to obtain a one-dimensional preliminary design parameter set maximizing the total static efficiency; and a hierarchical collaborative optimization framework is constructed to significantly improve the design efficiency and precision of the centripetal turbine, realize systematic approximation of global optimal design under the premise of guaranteeing physical authenticity, and provide a solution for high-performance turbine product development.