Abstract: The present invention relates to gas turbines and more particularly to a device for controlling the flow of cooling air through a flowpath in a turbine blade. The device can be inserted in the inlet opening of the blade flowpath and be retained therein. The device comprises a plug member for adjusting the flow of cooling air through the flowpath. The plug member comprises a retaining portion for retaining the plug member at the inlet opening of the flowpath and a blocking portion inserted within the flowpath for reducing the cross-sectional area of the inlet opening. Such a device is inexpensive and can be easily inserted in the inlet opening of a blade flowpath and retained therein.

Abstract: A method and device for decoupling combustor attenuation and pressure fluctuation from turbine attenuation and pressure fluctuation in a gas turbine engine. The engine has: a compressor; a combustor; and a turbine, that generate a flow of hot gas from the combustor to the turbine. An aerodynamic trip is disposed in at least one of; a combustor wall; and an inner shroud of the nozzle guide vane ring, and is adapted to emit jets of compressed air from cross flow ports into the flow of hot gas from the combustor. The air jets from the cross flow ports increase turbulence and equalize temperature distribution in addition to decoupling the attenuation and pressure fluctuations between the combustor and the turbine.

Abstract: A turbine vane for a gas turbine engine with an airfoil portion including a perimeter wall having first, second, third and fourth sets of cooling holes defined therethrough, including the holes numbered HA-1 to HA-13, HB-1 to HB-13, PA-1 to PA-9, and SA-1 to SA-3, respectively, and located such that a central axis thereof extends through the respective point 1 and point 2 having a nominal location in accordance with the X, Y, Z Cartesian coordinate values set forth in Table 3.

Abstract: The invention provides a novel sealing device and plenum assembly particularly adapted for forming a plenum inwardly of a circumferential stator blade array in a gas turbine engine. A Tangential On Board Injection plenum (T.O.B.I.) is defined by assembly of: a forward plate; a rearward plate; and a segmented inner shroud disposed between the forward and rearward plates. Releasable elongate fasteners are disposed in joints between a series of inner shroud segments with sealing sheaths laterally surrounding the fasteners. While clamping the shroud segments between the forward and rearward plates, the fasteners and sheaths simultaneously seal the joints and secure the inner shroud segments with forward and rearward plates in a rigid structural assembly.

Abstract: A seal spacer for a gas turbine engine to provide sealing between a casing wall and a stator vane of the engine. The seal spacer comprises a circumferential body having a slot defined therethrough and configured to receive the stator vane therein. The circumferential body is configured to be disposed between and space apart the engine casing wall from a head of the stator vane when the vane is inserted in the slot and the vane is inserted in the opening in the casing wall. The circumferential body includes a support positioned to be engage and provide a gap between the casing wall and the head when the seal spacer is installed therebetween, and an elastomeric seal portion positioned to engage and compressingly provide sealing the casing wall and the head when the seal spacer is installed therebetween to seal the gap.

Abstract: A stator blade assembly for a gas turbine engine has a leading edge portion with a passage communicating between a plenum and an air supply port of the outer shroud and an internal blade cooling channel communicating between passages and apertures adjacent the trailing edge of the blade. The plenum includes an impingement plate disposed a distance from the inner surface of blade platform to define an impingement cooling chamber within the plenum, and the plate includes impingement cooling apertures to direct cooling jets of air at the inner blade platform. An air flow restriction plate covers the inner end of the passage and controls the pressure and quantity of air delivered to the plenum via a compressed air metering aperture.

Abstract: An integrated strut and turbine vane nozzle (ISV) has inner and outer annular duct walls defining an annular flow passage therebetween. Circumferentially spaced-apart struts extend radially across the flow passage. Circumferentially spaced-apart vanes also extend radially across the flow passage and define a plurality of inter-vane passages. Each of the struts is integrated to an associated one of the vanes to form therewith an integrated strut-vane airfoil. The inter-vane passages on either side of the integrated strut-vane airfoil may be adjusted for aerodynamic considerations. The vanes may be made separately from the struts and manufactured such as to cater for potential misalignments between the parts.

Abstract: A turbine shroud segment has a body extending axially between a leading edge and a trailing edge and circumferentially between a first and a second lateral edge. A core cavity is defined in the body and extends axially from a front end adjacent the leading edge to a rear end adjacent to the trailing edge. A plurality of cooling inlets and outlets are respectively provided along the front end and the rear end of the core cavity. A crossover wall extends across the core cavity and defines a row of crossover holes configured to accelerate the flow of coolant directed into the core cavity via the cooling inlets. The crossover wall is positioned to accelerate the coolant flow at the beginning of the cooling scheme where the shroud segment is the most thermally solicited.

Abstract: A deflector for redirecting cooling air leakage to improve the efficiency of a gas turbine engine rotors. Adjacent rotor blades have a radially extending gap between each adjacent rotor blade and cooling air leaks into the gaps resulting in engine inefficiency. A deflector, disposed circumferentially on a forward face of the coverplate, deflects cooling air leakage, flowing from the running seal through the cooling air leakage path, away from the gaps between the blades, onto the outward surface of the blade platform, and into the gaspath at an acute angle relative to rearwardly axial flow. A forwardly extending flange deflects cooling air leakage forwardly around the blade platform lip.

Abstract: A fan assembly for a gas turbine engine is described which includes a fan and a leading edge assembly mounted to the fan. The leading edge assembly includes a plurality of leading edge extensions projecting from a central core and circumferentially spaced apart to align with leading edges of the fan blades. The leading edge extensions define cavities between the leading edges and the extensions. The cavities extend radially at least partially within the leading edge extensions, and receive heated pressurized air from the engine in operation. Elongated slots extend radially along a downstream edge of the leading edge extensions, and are defined axially between the downstream edge and the leading edges of the blades. The slots provide fluid flow communication between the cavities and at least pressure surfaces of the fan blades.

Abstract: There is provided an axial bearing load distribution system for a gas turbine engine of the type having a low pressure rotor supported by axial bearings. The system comprises a line having an inlet end positioned in a high-pressure compressor gas path downstream of any compressor stage provided with a variable geometry. The line is adapted to sense static pressure in the high-pressure compressor gas path. The line has an outlet end producing the static pressure. An air-tight pressure actuator is operatively connected to the outlet end and to one of the axial bearings to exert a force on the axial bearing proportionally to a pressure of the outlet end.

Abstract: A gas turbine engine has: a high-pressure turbine; a low-pressure turbine (LPT) having a LPT outer case and a LPT plenum extending around a central axis; a mid-turbine frame (MTF) assembly connecting the low-pressure turbine and the high-pressure turbine, the MTF assembly having: a MTF outer case defining an inlet for connection to a source of the cooling air, an intermediary plenum fluidly connected to the source of cooling air via the inlet, the intermediary plenum having a first air outlet and a second air outlet, and a MTF plenum disposed radially inwardly of the MTF outer case and fluidly connected to the source of the cooling air via the first air outlet of the intermediary plenum; and a cooling air conduit fluidly connecting the second air outlet of the intermediary plenum to the LPT plenum while bypassing the MTF plenum.

Abstract: A tangential on-board injector (TOBI) having: a body defining an annular passageway, discharge nozzles; a rotating component configured to be mounted for rotation relative to the body about an axis of rotation; a seal extending between the body and the rotating component; and flow passages circumferentially distributed about the axis of rotation and in fluid communication with the nozzles and the seal, the flow passages located upstream of the seal relative to a flow of the cooling air circulating toward the seal from the plurality of discharge nozzles, each of the flow passages extending along a respective passage axis, the passage axis of at least one of the flow passages having a tangential component at an outlet of the at least one of the flow passages that is different than a tangential component of an exit flow axis of at least one of the plurality of discharge nozzles.

Abstract: A turbine blade and assembly for use in a gas turbine engine are provided. The turbine blade has a root portion adapted for mounting to a rotor, and an airfoil portion extending from the root portion. A cooling air inlet duct is provided and adapted to communicate with the cooling air plenum when installed to the rotor. The air inlet duct has an inlet scoop portion adapted to extend into the cooling air plenum and an inlet scoop aperture oriented and adapted to capture cooling air from the cooling air plenum as a consequence of turbine rotation when the blade is mounted to the rotor. Scooped air is provided to a cooling air channel defined in an airfoil portion of the blade for the purpose of cooling the blade. The engine optionally includes a bearing gallery cooling air jacket in communication with the low pressure compressor stage for the purpose of cooling.

Abstract: A stator of a turbine section, has: vanes distributed around a central axis, a vane of the vanes extending along a spanwise axis and defining an internal passage; an insert received within the internal passage, the insert defining a cavity for receiving cooling air and defining impingement cooling apertures facing an inner face of the vane; a splitter plate secured within the cavity and being transverse to the spanwise axis, the splitter plate having a base secured to the insert and a tip protruding from the base; and a flow passage defined between the tip and the insert, the flow passage fluidly connecting a first section of the cavity to a second section of the cavity, the tip of the splitter plate secured to the insert at at least one location along a perimeter of the tip.

Abstract: A gas turbine engine rotor assembly such as a fan or compressor includes fan blades spaced axially from stator vanes inside an annular duct. A controlled air blowing system is secured to the annular duct and adapted to blow a hot air flow in a direction opposite to the rotation of the fan blades into the annular duct between the fan blades and the stator vanes, to impart an opposite rotational momentum with respect to the air compressed by the rotor blades, in a asymmetrical pattern to interfere with or destroy a symmetrical pattern of the blade rotation-wake of the compressed air, thereby reducing the strength or preventing the generation of the spinning mode so that the fan-stator interaction tones are significantly reduced.

Abstract: A turbine blade for a gas turbine engine has: an airfoil extending along a span from a base to a tip and along a chord from a leading edge to a trailing edge, the airfoil having a pressure side and a suction side, a tip pocket at the tip of the airfoil, the tip pocket at least partially surrounded by a peripheral tip wall defining a portion of the pressure and suction sides; at least one internal cooling passage in the airfoil and having at least one outlet communicating with the tip pocket; and a reinforcing bump located on the pressure side of the airfoil and protruding from a baseline surface of the peripheral tip wall to a bump end located into the tip pocket, the reinforcing bump overlapping a location where a curvature of a concave portion of the pressure side of the airfoil is maximal.

Abstract: An inlet guide vane assembly for an aircraft engine includes an array of inlet guide vanes having radially inner ends at a radially inner shroud, radially outer ends at a radially outer shroud, and airfoils extending therebetween. An internal passage extends radially through the airfoil from a vane air inlet at the radially inner end to a vane air outlet at the radially outer end. The vane air inlet is in fluid communication with an inner plenum, disposed radially inwardly of the inlet guide vane and in fluid communication with an anti-icing air source. The vane air outlet in fluid communication with an outer plenum, disposed radially outwardly of the inlet guide vane and having an exhaust port. A vane anti-icing pathway is defined in a radially-outward direction from the inner plenum, through the internal passage of the vane, and into the outer plenum.

Abstract: A method of operating a compressor of a gas turbine engine is described which includes directing a main airflow through tandem stator rows in a gaspath of the compressor, extracting a first portion of the main airflow from a first location proximate radially inner roots of stators of the first or second stator rows, extracting a second portion of the main airflow from a second location proximate the radially inner roots of the stators of the first or second stator rows, the second location being downstream of the first location relative to the main airflow, and re-injecting the combined extracted flow back into the main airflow at a third location. The third location is located upstream of the first and second locations, and is upstream of a leading edge of stators of the first stator row.

Abstract: A gas turbine engine blade tip clearance control system and method is described. An annular housing is formed about an engine casing to which an annular shroud segment assembly is secured and closely spaced about blade tips of a stage of blades. The annular housing forms an air passage means communicating with the casing for directing a cooling air stream to the casing. A thermally operable passive ring valve is formed by two overlapped metal ring segments having a dissimilar coefficient of thermal expansion selected whereby to produce a radial gap between the ring segments when the valve temperature reaches a predetermined value. The radial gap admits a cooling air flow into the housing for cooling the casing and its associated shroud segment assembly to control radial growth and thereby prevent blade tip pinching.