9629results about "Leakage prevention" patented technology

A turbine engine provides a spool supporting a turbine. The spool is arranged in a core nacelle and includes a thrust bearing. A fan is arranged upstream from the core nacelle and is coupled to the spool. A fan nacelle surrounds the fan and core nacelle and provides a bypass flow path that includes a fan nozzle exit area. A flow control device is adapted to effectively change the fan nozzle exit area. A controller is programmed to monitor the thrust bearing and command the flow control device in response to an undesired load on the thrust bearing. Effectively changing the fan nozzle exit area with the flow control device actively manages the bearing thrust load to desired levels.

A ceramic thermal barrier coating (TBC) (18) having first and second layers (20, 22), the second layer (22) having a lower thermal conductivity than the first layer for a given density. The second layer may be formed of a material with anisotropic crystal lattice structure. Voids (24) in at least the first layer (20) make the first layer less dense than the second layer. Grooves (28) are formed in the TBC (18) for thermal strain relief. The grooves may align with fluid streamlines over the TBC. Multiple layers (84, 86,88) may have respective sets of grooves (90), Preferred failure planes parallel to the coating surface (30) may be formed at different depths (A1, A2, A3) in the thickness of the TBC to stimulate generation of a fresh surface when a portion of the coating fails by spalling. A dense top layer (92) may provide environmental and erosion resistance.

A turbine ring assembly includes a ring support structure and a plurality of ring sectors, each including a single piece of ceramic matrix composite material. Each ring sector includes a first portion forming an annular base with an inside face defining an inside face of the turbine ring and an outside face from which there extends two tab-forming portions including ends that are engaged in housings in the ring support structure. The ring sectors present a section that is substantially π-shaped and the ends of the tabs are held without radial clearance by the ring support structure. The tabs can have a free length in meridian section that is not less than three times their mean width.

A seal housing is provided to substantially cover at least one duct wall of vane array duct of a gas turbine engine, and one example arrangement is employed in a mid-turbine frame. The arrangement provides improved sealing of the vane array duct through the provision of a plurality of cavities extending along the duct wall. The arrangement may also include insulation tubes to assist in sealing around load transfer spokes passing through the vane array.

A mid turbine frame is disposed between high and low pressure turbine assemblies. A cooling air system defined in the mid turbine frame of a gas turbine engine includes internal cavities for containing pressurized cooling air to cool the inter-turbine duct and the hollow struts, and discharges the used cooling air to further cool respective high and low pressure turbine assemblies.

A material system (60) contains close packed hollow shapes (50, 70) having a dense wall structure (52, 66), which are bonded together and which may contain a matrix binder material (56) between the shapes, where the system has a stable porosity, and is abradable and thermally stable at temperatures up to possibly 1700° C., where such systems are useful in turbine apparatus.

A turbine shroud assembly includes forward and aft hangers, an axisymmetric plenum assembly, ceramic shroud segments, ceramic spacers, and forward and aft rope seals. The plenum assembly supplies impingement cooling to the shroud and the hangers. The impingement cooling to the forward and aft hangers is controlled independently to improve blade tip clearance. The rope seals are radially inward from the hangers and reduce cooling flow leakage. The turbine shroud assembly can operate in a higher temperature environment using less cooling flow than the prior art.

A turbine engine has a first disk and a second disk, each extending radially from an inner aperture to an outer periphery. A coupling, transmits a torque and a longitudinal compressive force between the first and second disks. The coupling has first means for transmitting a majority of the torque and a majority of the force and second means, radially outboard of the first means, for vibration stabilizing.

An object of the present invention is to provide a ring segment of a gas turbine in which the temperature is maintained low, damage due to high temperature oxidation is prevented, and high temperature deformation is prevented. In order to achieve the object, the present invention provides a ring segment of a gas turbine which comprises a blade ring, a main shaft and moving blades comprising a plurality of individual units which define an annular form by being arranged around the peripheral direction of the main shaft, and disposed so that its inner peripheral surface is maintained at a constant distance from the tips of the moving blades, wherein grooves which extend along the axial direction of the main shaft of the turbine are formed upon of the individual units so as mutually to confront one another; a seal plate which is inserted into each mutually confronting pair of the grooves so as to connect together the adjacent pair of individual units; and contact surfaces which are formed at positions more radially inward than the seal plates, which extend in the axial direction and the peripheral direction and which mutually contact one another.

The coating of abradable material presents cavities opening out into the surface of the coating, the cavities extending over a fraction of the thickness of the coating and being defined by walls which form a plurality of continuous paths at the inside surface of the annular portion between the axial end faces thereof, and which form an angle of not less than 5° relative to the general direction of end portions of blades that might come into contact with the abradable material.

An electrically controlled turbocharger has a motor mounted on a shaft in a motor housing between a turbine and compressor. Oil is sprayed onto the motor stator to cool the stator. An oil gallery is disposed above the stator to receive lubricating oil and contains apertures that perform as jets to allow oil to be sprayed directly on the motor stator. A coolant jacket is formed in the turbocharger housing between the turbine and the motor to allow liquid coolant to circulate therein and dissipate heat from the turbine end prior to reaching the motor components. Other embodiments provide for a stator component to be submerged in flowing cooling oil.

A cooling fan for an engine in a vehicle. Ordinarily, a fan rotates within a shroud, which surrounds the fan. Leakage can occur between the tips of the fan blades and the shroud, wherein fan exhaust moves forward, and then passes through the fan again. The invention reduces leakage by placing a surface downstream of the fan. The surface employs the Coanda Effect, to urge fan exhaust to continue in the downstream direction, and not move forward as leakage air.

At least one shroud segment, floating axially independently of adjacent turbine engine shroud assembly members, includes a segment body comprising a radially outer surface and a radially outwardly projecting segment support that includes an axial support wall surface therein. The assembly includes a shroud hanger in axial juxtaposition with the segment support, and at least one axial support projection from the shroud hanger into the segment support at the support wall surface. The support projection supports the shroud segment releasably at the support wall surface sufficiently to enable relative axial movement of the shroud segment on the support projection independently of the shroud hanger and adjacent engine members.

Methods for repairing and manufacturing a gas turbine blade, and the gas turbine blade repaired and manufactured with such methods are presented with, for example, the repair method comprising providing a gas turbine blade, the blade comprising a blade tip and a blade body; removing at least one portion of the blade tip; providing at least one freestanding tip insert; and disposing the at least one tip insert onto the gas turbine blade body such that the at least one tip insert replaces the at least one removed portion of the blade tip.