87 results about "Tip clearance" patented technology

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.

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.

The invention discloses a method for measuring a blade tip gap of a metal blade of an aero-engine by an eddy current method, and the method comprises the steps: calculating a mutual inductance distance between a probe coil and a blade top plane through an eddy current displacement sensor, a voltage / current-equivalent inductance quality factor calibration curve, an equivalent inductance quality factor and a mutual inductance distance relational expression between the probe coil and the blade top plane; the effective mutual inductance area is corrected by considering the included angle between the axis of the probe coil and the local normal of the blade top plane and the axial movement of the blade, a calculation formula of the relative influence factors under the actual temperature is established, and the mutual inductance distance between the probe coil and the blade top plane is corrected by considering the installation angle deviation existing in the eddy current displacement sensor. And finally calculating the mutual inductance distance between the blade top surface and the probe coil when the included angle between the central axis of the probe coil and the local normal of the blade top is 0, namely the blade tip gap. The method can achieve the detection of the dynamic tip clearance of the aero-engine, and is high in detection precision.

The application discloses a blade tip structure and blade with improved cooling efficiency, comprising a groove formed in a tip clearance leakage flow area, the inner wall of the groove is inclined to the mean camber line of the blade, the outlet of the cooling hole of the blade is located on the inner wall of the groove, the flow direction of the cooling working medium output by the cooling hole is the same as the flow direction of the leakage flow, the dovetail groove and the cooling hole are combined, the dovetail groove suppresses the formation of the pressure side corner vortex and reduces the heat exchange coefficient on the near pressure surface side; the cooling hole directly blows out the cooling gas to the blade tip high heat exchange area for impact cooling, the cooling gas flows to the suction surface side under the action of the pressure difference after the impact cooling, the blade tip temperature distribution is improved, the cooling gas utilization rate is improved, and the blade tip temperature distribution is uniform; on the other hand, the high-speed cooling jet forms an approximate circumferential rib structure, plays a role of aerodynamic sealing, suppresses the development of the leakage flow to the downstream, reduces the blade tip leakage amount, improves the gas-thermal characteristics of the blade tip area, prolongs the service life of the high-pressure turbine blade, and guarantees the safe and effective operation of the gas turbine.

The invention provides a real-time prediction method and system for a turbine blade tip clearance, a storage medium and a computer program product. The real-time prediction method for the turbine blade tip clearance comprises the steps that the overall performance parameters of an engine are obtained; establishing a component heat transfer model of the engine, and inputting the overall performance parameters of the engine into the component heat transfer model to obtain a real-time average temperature of a component; establishing a part deformation model of the engine, and inputting the real-time average temperature of the part into the part deformation model to obtain part deformation; and the real-time turbine blade tip clearance is obtained according to the part deformation. According to the real-time prediction method, the turbine blade tip clearance can be rapidly and accurately predicted.

The application discloses an air suspension centrifugal blower capable of avoiding large-amplitude airflow oscillation, comprising a centrifugal blower box, a high-speed permanent magnet motor arranged in the box, an impeller connected with an output shaft of the high-speed permanent magnet motor and located in a volute, and a suction chamber connected with the middle part of the volute; an adjusting mechanism for adjusting a blade tip clearance is arranged on a wheel shaft of the impeller. The application further discloses a processing technology of the air suspension centrifugal blower capable of avoiding large-amplitude airflow oscillation, wherein, after dynamic balance testing of the wheel shaft and the impeller, a radial bearing and an axial bearing are assembled; the impeller is assembled into the volute and installed on the high-speed permanent magnet motor; soundproof cotton is installed on the inner wall of the box, and the high-speed permanent magnet motor is installed into the box. The application can effectively control the blade tip clearance to prevent surge, especially when the flow is small, the blade tip clearance is reduced, and the back pressure at the outlet of the impeller is increased; the flow entering the blade of the suction chamber is away from the surge zone, and the surge is effectively prevented.

This invention relates to a low-noise axial flow blade with dual-stage serrations. It pertains to the field of axial flow fan technology. The blade trailing edge is divided into a lower-middle region (I) and an upper region near the blade tip (II) along the blade height. Both regions have serrations on their trailing edges, forming a dual-stage noise reduction structure. The trailing edge of region I is fitted using a Mie-squared function to match gradually sized serrations; the trailing edge of region II is fitted using a linear square root mixed function to match equidistant serrations. This invention can simultaneously and specifically suppress two core noise sources: trailing edge shedding vortices and tip clearance leakage vortices, significantly reducing fan aerodynamic noise while maintaining blade aerodynamic efficiency. It can be directly adapted to existing conventional fan installation structures, has a mature molding process, and is suitable for low-pressure, high-volume applications such as HVAC and industrial ventilation.

The invention relates to a gas turbine engine, a turbine tip clearance control method and system thereof and a medium. The control method comprises the steps that S1, the steady-state wall temperature of a casing, the total pressure of an ACC pipeline inlet, the total temperature of the ACC pipeline inlet, the environment pressure of an engine, the total temperature of a gas compressor outlet and the flow of the gas compressor inlet when the ACC is closed are obtained; s2, according to the data obtained in the step S1, the temperature drop and the cooling flow of the casing are obtained; and S3, according to the data obtained in the S2, the ACC valve angle is obtained.

The present application relates to fan mechanical technology field, especially to a kind of moving blade assembly, axial flow fan and range hood, including shell and the hub being set in the shell;Multiple blades, with the rotary axis of the hub as center symmetrically arranged, the top of the blade and the inner wall of the shell are spaced, the blade has suction surface and pressure surface;First shoulder, protruding on the suction surface, along the flow direction of airflow, the first shoulder is inclined and extended upwards on the suction surface, so that the first shoulder and the inner wall of the shell form tapered airflow passage;And, second shoulder, protruding on the pressure surface, along the flow direction, the second shoulder is inclined and extended downwards on the pressure surface, so that the second shoulder and the inner wall of the shell form gradually expanded airflow passage.The moving blade assembly structure can reduce the leakage flow in airflow passage without changing the blade tip clearance.

The invention provides a dynamic blade tip clearance control method for an axial flow fan. The dynamic blade tip clearance control method comprises the following steps that S1, a multi-physical field coupling analysis model of an impeller-frame system is constructed; s2, on the basis of the multi-physics field coupling analysis model, simulating calculation is conducted on the maximum radial displacement delta Rrootor of the impeller and the minimum radial displacement delta Rhousing of the inner circle of the frame, and the dynamic envelope range of the outer edge of the impeller and the dynamic envelope range of the inner edge of the frame under the design working condition are obtained; s3, calculating a minimum safety gap required between the impeller and the frame; and S4, according to the minimum safety clearance delta min obtained through calculation, fan mechanical design is guided, and it is ensured that the actual assembly blade tip clearance is larger than or equal to delta min. According to the method, the minimum safety clearance is accurately calculated, fan design and manufacturing are guided, and minimization and optimization of the blade tip clearance are achieved on the premise that operation interference is avoided.

The application discloses a turbine tip clearance passive matching method based on engine demand, belongs to the technical field of aero-engines, and comprises the following steps: determining a minimum allowable working clearance; determining a typical tip clearance analysis load spectrum; performing thermal analysis and deformation analysis on a rotor and a stator; performing tip clearance change rule analysis according to the deformation results of the rotor and the stator to obtain a tip clearance change rule; optimizing tip clearance matching according to different engine demands, the casing material, the structure and the cooling form; determining an initial cold-state clearance according to a narrow-point clearance and the minimum allowable working clearance; and performing iterative calculation to determine a final tip clearance matching scheme. The turbine rotor deformation and the turbine stator deformation are matched in the application, so that the turbine tip clearance can be as small as possible under the operating states that are paid attention to, that is, the turbine tip clearance is as small as possible under the premise of ensuring the safety of the engine, the turbine efficiency is improved, and the specific fuel consumption is reduced.

The utility model belongs to aero -engine turbine blade technical field provides an aero -engine turbine blade tip gap adjustable sealing device, including the casing body, the first installation groove is set up in casing body inboard wall, the fixed ring is fixedly connected with the casing body outside wall, the movable sleeve of casing body outside wall has the actuating ring, the inside of fixed ring is uniformly provided with a plurality of adjusting assembly, the inside of actuating ring is uniformly provided with a plurality of locking assembly, the utility model discloses through the cooperation of actuating ring, rocker, adjusting assembly, can adjust tip gap according to turbine front temperature change, guarantee turbine to keep reasonable tip gap under different operating conditions, improve turbine efficiency, reduce fuel consumption, the setting of locking assembly makes actuating ring realize self -locking after adjusting, avoids the change of gap due to actuating ring rotation, effectively prevents the collision of tip and turbine outer ring, improves the safety and reliability of engine work.