Abstract: A method of manufacturing a compressor stator having: a first stator blade with a first leading edge and a first trailing edge; a second stator blade disposed a circumferential distance from the first stator blade, the second stator blade having a second leading edge disposed an axial distance from the first leading edge and a second trailing edge disposed an axial distance from the first trailing edge; the method comprising: using additive manufacturing to deposit and fuse together progressive layers of metal material commencing at a substrate to form the first stator blade, the second stator blade, at least one intermediate support structure disposed between the first stator blade and the second stator blade, and at least one primary support structure disposed between the substrate and at least one of: the first stator blade; and the second stator blade; and removing the primary support structure and the intermediate support structure.

Abstract: A fan blade anti-icing system comprises a fan hub and a fan blade extending radially outwardly from the fan hub. The fan blade has a base and an airfoil extending radially outwardly from the fan base. The airfoil having a leading edge, a trailing edge, a convex side surface between the leading and trailing edge and a concave side surface between the leading and trailing edge. The fan blade further has a radial passage extending from a blade air inlet in the blade base in communication with a source of heated air, and a rearwardly directed passage in communication with the radial passage and having a blade air outlet forward of the trailing edge and oriented tangentially to the convex side surface or concave side surface of the airfoil.

Abstract: A gas turbine engine airfoil has a hollow airfoil section extending chordwise between a leading edge and a trailing edge. The airfoil has a leading edge cooling passage and a separate serpentine passage for cooling a remaining portion of the airfoil. The serpentine passage has at least three segment serially interconnected in fluid flow communication. The leading edge cooling passage and the serpentine cooling passage have separate coolant inlets. The coolant inlet of the serpentine passage comprises a primary inlet branch connected in fluid flow communication with a first one of the segments of the serpentine passage and a secondary inlet branch connected in flow communication with a last one of the segments, thereby providing for a portion of the flow passing through the coolant inlet of the serpentine passage to be directly fed into the last segment of the serpentine passage.

Abstract: A method and device for optically measuring throat area in a vane ring for a gas turbine engine, the method according to one aspect of the invention following the steps of: placing the vane ring on a fixture with the periphery of each throat within an optical measuring field of view; positioning a primary radiation source to cast an area of shadow on the vane ring delineating the throat as a dark area surrounded by an area of reflectance; capturing an image of the dark area with a radiation detector; analyzing the image to acquire a dimensional data of the dark area, proportional to the absolute value of throat dimensions; progressively capturing and analyzing images from each of the individual throats; then processing and calibrating the dimensional data of each image to account for scaling and viewing direction (perspective distortion) to acquire an absolute value for the composite throat area of the vane ring.

Abstract: A tangential on-board injector (TOBI), comprising: a body defining an annular passageway to receive cooling air, the TOBI defining a plurality of discharge nozzles; a rotating component mounted for rotation relative to the body about an axis of rotation; a seal extending between the body and the rotating component; a plurality of vanes circumferentially distributed about the axis of rotation and located downstream of the plurality of discharge nozzles relative to a flow of the cooling air circulating toward the seal from the plurality of discharge nozzles and upstream of the seal; and flow passages defined between the plurality of vanes, a flow passage of the flow passages extending along a passage axis, the passage axis having a tangential component at an outlet of the flow passage that is different than a tangential component of an exit flow axis of a nozzle of the plurality of discharge nozzles.

Abstract: A gas turbine engine includes an annular casing and a plurality of shroud segments forming an annular shroud. The annular shroud forms with the annular casing an annular cavity therebetween. The annular cavity includes an inlet and an outlet. An annular ring assembly is disposed in the annular cavity between the casing and the shroud and cooperating therewith to provide a first annular chamber and a second annular chamber. The annular ring assembly and a first portion of the shroud form the first annular chamber. The annular ring assembly and a second portion of the shroud form the second annular chamber. The annular ring assembly forms an intermediate annular chamber disposed between the first annular chamber and the second annular chamber. A flow path for coolant air is sequentially defined through the inlet, the first annular chamber, the intermediate annular chamber, the second annular chamber and the outlet.

Abstract: An internally cooled turbine vane for a gas turbine engine has coolant flow channels between the interior walls of the vane and an insert, where the channels serve to convey a portion of the cooling air flow from a pressure side chamber to a suction side chamber. The turbine vane defines a radially extending passage with a dividing wall defining a front section and a rear section; the rear section having interior walls spaced apart from an insert to define the pressure side chamber and the suction side chamber. The insert may receive cooling air and conveys the cooling air into the pressure side chamber and the suction side chamber. A front surface of the insert or a rear surface of the dividing wall may have a clearance gap and an air flow channel communicating between the pressure side chamber and the suction side chamber.

Abstract: A rotor containment structure for gas turbine engine comprises an inner containment layer having a single integral body defining an annular structure about the rotor, and having a support on the inner surface of the inner containment layer for at least one shroud segment. An outer containment layer provides containment strength to contain blade fragments and defining an annular structure about the inner containment layer. Air passage through the outer containment layer for air to pass from an exterior of the outer containment layer to an interior of the outer containment layer; and the inner containment layer being connected at a first end to the outer containment layer with a gap defined between the inner surface of the outer containment layer and the outer surface of the inner containment layer, the gap being in fluid communication with the air passage such that air flows through the gap, beyond a free second end of the inner containment layer.

Abstract: A tangential on board injector with auxiliary supply of further cooled compressed air, from an external heat exchanger or air cooled bearing gallery for example, serves to reduce the volume of cooling air directed tangentially toward a cooled rotor of a gas turbine engine. The tangential on board injector has an array of injector blades between two injector walls defining circumferential main flow nozzles for directing a main compressed air flow tangentially. Each blade has an interior chamber in flow communication with a source of auxiliary compressed air with at least one bore extending between the chamber and an exterior surface of the blade. The bores eject further cooled air from the heat exchanger and merge with the primary compressed air flowing through the injector nozzles. The bores may also produce a cooling film of air that reduces drag over the injector blades.

Abstract: A Tangential On-Board Injection (TOBI) nozzle or a Radial On-Board Injection (ROBI) nozzle with adjustable flow area in a gas turbine engine for directing a cooling air flow into a rotor assembly of the turbine is disclosed. The TOBI or ROBI nozzle is a unitary cast structure including a plurality of tangentially angled cooling air passages, a number of the passages being cast-closed. After a flow test is conducted using the TOBI or ROBI nozzle, one or more of the cast-closed passages may be selectively machined open if the mass flow of the TOBI or ROBI nozzle does not satisfy the design point. Therefore, the structure of the TOBI or ROBI nozzle is easily fabricated using a casting process with low manufacturing costs while the optimum cooling air flow is achieved by adjusting the total flow area of the TOBI or ROBI nozzle before it is installed to the gas turbine engine.