16 results about "Tip clearance" patented technology

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.

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.

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 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.

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 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.

An operating method is provided during which a command is received to increase thrust generated by a propulsor rotor from a first thrust level to a second thrust level. The propulsor rotor is operatively coupled to an engine core and an electric machine. The engine core includes a flowpath, a compressor section, a combustor section and a turbine section. The engine core is operated in a transient state to increase power output from the engine core to the propulsor rotor from a first power level to a second power level in response to the command. The electric machine is operated to boost the power output from the engine core to the propulsor rotor while the engine core is operating in the transient state. A clearance control system for the engine core is operated based on the operation of the electric machine.

This invention relates to the field of aero-engine technology, and discloses an adjustable dual-blade tip clearance simulation test device and method for rotor elastic line reconstruction and fault identification tests. The invention employs a dual independent drive assembly design. During rotor rotation, the axial displacement of the first and second auxiliary bearings is controlled by the first and second drive assemblies, respectively. The axial movement of the first and second auxiliary bearings is converted into radial extension and retraction of the blades via a connecting rod structure, allowing both adjustable blades to extend or retract simultaneously, achieving synchronous adjustment of the tip clearance of both blades. Alternatively, the tip clearance of a single adjustable blade can be adjusted by extending or retracting it individually. This invention can meet the dynamic adjustment requirements of tip clearance in rotor elastic line reconstruction and fault identification tests, and can support rotor dynamics research in rotating turbomachinery.

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 purpose of this invention is to provide a turbine casing-blade tip cross-coupling structure for reducing axial leakage flow, belonging to the field of gas turbine blades. Its specific structure is as follows: spaced-apart blade tip edge ribs are provided at the top of the blade body, and cross grooves are formed between adjacent blade tip edge ribs; casing ribs are provided at the bottom of the casing, and the casing ribs are installed in their corresponding blade tip edge rib cross grooves; a blade tip clearance exists between the casing and the blade body; and a blade tip groove is provided inside the blade body below the blade tip clearance. This invention can significantly reduce the axial leakage velocity at the blade tip and improve aerodynamic efficiency; the efficient blocking effect of the cross-coupling structure reduces the material required for traditional grooved blade tips, achieving a lightweight blade tip design and reducing the centrifugal load on rotating components; the design parameters of this invention are highly adjustable and can be flexibly optimized according to different operating conditions, possessing wide applicability.

This application relates to the field of air conditioning technology, and in particular to a ducted air conditioning unit. The ducted air conditioning unit proposed in this application includes multiple fans and a drive unit. The drive unit is disposed between adjacent fans and includes a motor and a connecting shaft. The motor drives the fans of the multiple fans to rotate through the connecting shaft. By constructing a coordinated matching relationship between the total length of the volute, the fan spacing, and the motor size, and limiting the ratio of the total volute length W to the total spacing S, the volute opening W / S is limited to the optimal range of 0.55~0.675. This allows the fan blades to form a reasonable tip clearance with the inner wall of the casing, effectively suppressing high-frequency aerodynamic noise generated by tip vortices, while ensuring smooth airflow within the casing. This reduces airflow turbulence and increased vortex noise caused by excessive clearance, and prevents airflow interference, increased structural vibration, and structural interference risks caused by insufficient clearance. This achieves a comprehensive performance improvement of low noise, large air volume, and high stability from the structural layout source.

This invention relates to a high-efficiency cooling system for supersonic turbine blade tips based on shock wave interference effects. Specifically, it involves setting film cooling holes of different sizes in the leading and trailing edge regions of the turbine blade. Through numerical simulation, the wave system structure of the blade tip clearance is accurately captured, clarifying the distribution of high-pressure / low-pressure regions formed by shock wave interference. The relationship between shock waves and tip heat transfer and flow field is precisely analyzed, as well as the flow field distribution within the clearance, which differs from conventional subsonic blade tip clearances, due to the complex wave system structure of shock waves and expansion waves. By arranging film cooling holes of different diameters and spacings close to the pressure side in the leading and trailing edge regions, the non-uniform spatial layout of the cooling holes is optimized, greatly improving cooling efficiency and targeting. This solves the cooling failure problem caused by neglecting the influence of shock wave interference, achieving precise matching between flow field characteristics and cooling structure design, resulting in greater targeting.

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.

An aircraft engine having a turbine assembly including a bearing housing having a central axis, a first-stage vane ring and a second-stage vane ring. The first-stage vane ring is positioned in concentric relation with the bearing housing via a first circumferential array of lugs and slots between the bearing housing and the first-stage vane ring. The second-stage vane ring is positioned in concentric relation to the first-stage vane ring via a second circumferential array of lugs and slots between the first-stage vane ring and the second-stage vane ring.