Patent application number | Description | Published |
20110016877 | CONTINUOUS SLOT IN SHROUD - A compressor shroud has a continuous slot defined in a gaspath side surface thereof for communicating through the body of the shroud. Spaced-apart structural bridges span a rear portion of the slot in the shroud to maintain structural rigidity of the shroud. The method of forming the continuous slot is also described. | 01-27-2011 |
20130236319 | AIRFOIL FOR GAS TURBINE ENGINE - An airfoil for a rotor blade or a stator vane of a gas turbine engine having a tip extension extending one or both of the suction and pressure side surfaces adjacent the tip. The tip extension defines a thickness at the tip larger than the true thickness of the airfoil. The tip extension includes a side transitional surface substantially defined along a curve extending tangentially from the corresponding side surface. The tip extension defines a gradual increase in the thickness over a portion of the airfoil having a radial dimension corresponding to at least 2 times the true thickness. | 09-12-2013 |
20130287542 | TWISTED VARIABLE INLET GUIDE VANE - A compressor for a gas turbine engine with variable inlet guide vanes each defining an airfoil portion twisted such that at each location of the airfoil portion along the pivot axis, an angle is defined between a respective chord extending between the leading and trailing edges and a same reference plane containing the pivot axis and extending radially with respect to the compressor. The angle, which is measured along a direction of rotation of the rotor, varies from a minimum value near the hub side wall to a maximum value near the shroud side wall. A method of reducing vortex whistle in a radial inlet of a compressor is also provided. | 10-31-2013 |
20140271108 | COMPRESSOR BLEED SELF-RECIRCULATING SYSTEM - A compressor for a gas turbine engine having a bleed air recirculation system includes a plurality of bleed holes extending through the shroud at a first axial location thereon substantially adjacent the blade tips. The bleed holes have a closed outer perimeter along their complete length. An annular bleed cavity surrounds the shroud and is in communication with outlet openings of the bleed holes. The bleed holes provide communication between said main gas flow passage and the bleed cavity. The bleed cavity includes exit passages having outlets disposed in said shroud at a second axial location which is upstream of both the first axial location and the leading edge of the blades of the rotor. Bleed air is passively bled from the main gas flow passage via the bleed holes, recirculated through the bleed cavity and re-injected back into the main gas flow passage at the second axial location. | 09-18-2014 |
20140271158 | COMPRESSOR STATOR - Compressor stators ( | 09-18-2014 |
20140286746 | COMPRESSOR SHROUD REVERSE BLEED HOLES - A gas turbine engine compressor includes a rotor defining a central axis of rotation and a plurality of blades which project into an annular compressor gas flow passage, and a shroud circumferentially surrounding the rotor and having a radially inner surface adjacent to the blade tips. Bleed holes extend through the shroud adjacent the blade tips, each of the bleed holes having an inlet end disposed in the shroud radially inner surface and an outlet end disposed in a shroud radially outer surface. Bleed air removed from the annular gas flow passage flows through the bleed holes from the inlet to the outlet ends. The outlet end of each bleed hole is located circumferentially upstream of the inlet end relative to a direction of rotational flow in the annular gas flow passage driven by a direction of rotation of the rotor. | 09-25-2014 |
20140356159 | LOW HUB-TO-TIP RATIO FAN FOR A TURBOFAN GAS TURBINE ENGINE - A fan for a turbofan gas turbine engine, the fan comprising a rotor hub and a plurality of radially extending fan blades integral with the hub to form an integrally bladed rotor. Each fan blade defines a leading edge. A hub radius (R | 12-04-2014 |
20150219115 | Blade for Axial Compressor Rotor - A blade for a compressor rotor comprises a blade root at a first end of the blade extending from a rotor body of the compressor, and a blade tip at a second end of the blade. The first end and the second end are disposed on a span of the blade. The blade root and the blade tip define a face of the blade between them. The blade tip comprises a first extremity and a second extremity. The blade tip is disposed such that a first segment of the blade tip defines a positive dihedral angle with the face of the blade. The first segment extends from a first point in the blade tip to the first extremity of the blade tip. A second segment of the blade tip defines a negative dihedral angle with the face of the blade. The second segment extends from a second point in the blade tip to the second extremity of the blade tip. | 08-06-2015 |