Abstract: A gas turbine shaft assembly comprises first and second axially mating shafts. The first shaft has a circumferential array of splines configured to mesh with a corresponding circumferential array of splines on the second shaft. The first shaft defines at least one axially extending slot aligned with one of the splines on the first shaft. The second shaft has at least one radial projection aligned with one of the splines on the second shaft. The engagement of the radial projection with the slot provides angular correspondence between the splines of the first and the second shafts prior to axial engagement of the splines.

Abstract: A gas turbine engine has an in-line mounted accessory gear box (AGB) and an accessory drivingly connected to the AGB, the accessory being oriented transversally to the engine centerline.

Abstract: A gas turbine engine has a first spool having a low pressure compressor section disposed forward of an air inlet along a direction of travel of the engine, and a low pressure turbine section disposed forward of the low pressure compressor section and drivingly engaged thereto. A second spool has a high pressure compressor section disposed forward of the low pressure compressor section, and a high pressure turbine section disposed forward of the high pressure compressor section and drivingly engaged thereto. The high pressure turbine section is disposed aft of the low pressure turbine section. An output drive shaft drivingly engages the low pressure turbine section and extends forwardly therefrom to drive a rotatable load. A method of operating a gas turbine engine is also discussed.

Abstract: A multi-spool gas turbine engine comprises a low pressure (LP) spool and a high pressure (HP) spool independently rotatable about a central axis extending through an accessory gear box (AGB). The LP spool has an LP compressor, which is axially positioned between the HP compressor of the HP spool and the AGB. A tower shaft drivingly connects the HP spool to the AGB.

Abstract: A method of operating a twin engine helicopter power plant, the power plant comprising: two turboshaft engines each having an engine shaft with a turbine at a distal end and a one-way clutch at a proximal end; a gear box having an input driven by the one way clutch of each engine and an output driving a helicopter rotor; a bypass clutch disposed between the proximal end of each engine shaft and the input of the gear box; and power plant management system controls for activating the bypass clutch; the method comprising: detecting when a rotary speed of an associated engine shaft is less than a rotary speed of the gear box input; activating the bypass clutch to drive the associated engine shaft using the rotation of the gear box input; and starting an associated engine by injecting fuel when the bypass clutch is activated.

Abstract: A bearing housing comprises a body having a partition wall delimiting the bearing housing from an environment and defining a bearing housing interior cavity configured to receive an oil feed. An inlet bore is in the partition wall in fluid communication with the bearing housing interior cavity. The inlet bore is configured to receive an end of an oil tube, the inlet bore comprising a contact surface configured for contacting the oil tube. A seal is between the inlet bore and the oil tube. An oil recuperation passage is defined in the partition wall in fluid communication with the inlet bore between the seal and at least a portion of the contact surface, the oil recuperation passage being in fluid communication with an oil recuperating cavity.

Abstract: A bearing housing comprises a body having a partition wall delimiting the bearing housing from an environment and defining a bearing housing interior cavity configured to receive an oil feed. An inlet bore is in the partition wall in fluid communication with the bearing housing interior cavity. The inlet bore is configured to receive an end of an oil tube, the inlet bore comprising a contact surface configured for contacting the oil tube. A seal is between the inlet bore and the oil tube. An oil recuperation passage is defined in the partition wall in fluid communication with the inlet bore between the seal and at least a portion of the contact surface, the oil recuperation passage being in fluid communication with an oil recuperating cavity.

Abstract: A turbine shroud segment is metal injection molded (MIM) about a low melting point material insert. The low melting point material is dissolved using heat during the heat treatment cycle required for the MIM material, thereby leaving internal cooling passages in the MIM shroud segment without extra manufacturing operation.

Abstract: A method, system and computer program product for distinguishing superimposed links to a node in a drawing area. Links connected to a side of a node involved in an editing interaction are identified. The order of the identified links to be connected to the side of the node is determined by sorting the identified links based on the coordinates of the links' bend point and/or opposite end point. The connection points of these identified links are repositioned based on this determined order. By only repositioning the connection points to the node involved in the editing interaction instead of rerouting all of the links in the diagram area, the processing time is improved. Furthermore, by repositioning the connection points that were superimposed so that they are spaced apart based on the coordinates of the links' bend point and/or opposite end point, it is easier to distinguish between the different links.

Abstract: A turbine shroud segment is metal injection molded (MIM) about a low melting point material insert. The low melting point material is dissolved using heat during the heat treatment cycle required for the MIM material, thereby leaving internal cooling passages in the MIM shroud segment without extra manufacturing operation.

Abstract: An exhaust mixer for a gas turbine engine includes an annular wall having upstream end adapted to be fastened to an engine case and a downstream end forming a plurality of inner and outer mixer lobes. A support member interconnects at least a number of the inner lobes, and includes a circumferentially extending stiffener ring located radially inwardly from the inner lobes and a series of circumferentially spaced apart mixer struts radially extending from the inner lobes to the stiffener ring. The mixer struts have a radial length at least equal to a width of a main gas path defined between the inner lobes and the exhaust cone such that the mixer struts extend entirely through the main gas path. The stiffener ring being fixed solely to the mixer struts such as to float with respect to the exhaust cone and permit relative movement therebetween.

Abstract: An exhaust mixer for a gas turbine engine includes an annular wall having upstream end adapted to be fastened to an engine case and a downstream end forming a plurality of inner and outer mixer lobes. A support member interconnects at least a number of the inner lobes, and includes a circumferentially extending stiffener ring located radially inwardly from the inner lobes and a series of circumferentially spaced apart mixer struts radially extending from the inner lobes to the stiffener ring. The mixer struts have a radial length at least equal to a width of a main gas path defined between the inner lobes and the exhaust cone such that the mixer struts extend entirely through the main gas path. The stiffener ring being fixed solely to the mixer struts such as to float with respect to the exhaust cone and permit relative movement therebetween.

Abstract: A thermal expansion joint suited for metal sheet component assemblies and method of connection is described. The expansion joint connection is formed by overlapping edge portions of components which are subject to different temperatures. Slot apertures are formed in the edge portion of a first component at predetermined locations and hole apertures are formed in the edge portion of the second component and disposed for confronting alignment with the slot apertures. A connecting sleeve is disposed in each hole aperture and projects through the hole and associated slot and a retention washer is secured over an end of the connecting sleeve on the other side of the second component. A fastener extends through the retention washer and the connecting sleeve to clampingly interconnect the overlapped edge portions of the first and second components.

Abstract: A mid-turbine frame is disposed between high and low pressure turbine assemblies. A secondary air system is defined in the mid-turbine frame (MTF) to provide cooling to the turbine section of the engine. The secondary air system may be used to cool and pressurize seals to assist with oil retention in bearing cavities. The temperature gain of the secondary air may be reduced by flowing the secondary air through one or more external lines and then generally radially inwardly through air passages defined in the MTF.

Abstract: A gas turbine engine case has a metallic composite boss mounted thereto. The boss may have a main metal injection molded (MIM) part, and a separately formed flange part integrated to the main MIM part and projecting laterally outwardly therefrom. The MIM part and the flange part may be made of different materials. The material of the flange part is selected to provide a suitable mounting interface with the gas turbine engine case.

Abstract: A turbine vane of a gas turbine engine is provided with a hollow core in the leading edge of the outer platform thereof. The core is interconnected with the leading edge core of the airfoil whereby to create a cooling air stream having a dual purpose and cooling both the leading edge of the outer platform and of the airfoil and thereby reducing cooling air consumption. The cooling air enters the core of the outer platform through an inlet port and exits through cooling holes provided in the leading edge of the airfoil.

Abstract: A method of manufacturing a turbine shroud segment for a gas turbine engine comprises: metal injection molding (MIM) a shroud segment body with a groove defined in a first lateral side thereof and with a flow restrictor projecting integrally from an opposite second lateral side thereof. The groove is oversized relative to the flow restrictor to provide for a clearance fit between the flow restrictor and the groove of adjacent turbine shroud segments when assembled together in a ring formation. The shroud segment body with the integrated flow restrictor are then subjected to debinding and sintering operations.

Abstract: A stiffener rib for a non-cooled, hollow core airfoil of a turbine blade is described. The stiffener rib is constituted by an elongated solid metal piece dimensioned for securement in a hollow region of a core portion of the airfoil. The stiffener rib is shaped and oriented upwards from a leading edge of the airfoil at a predetermined calculated angle to minimize mode 2 deformation of the blade at a trailing edge thereof while improving the rigidity of the turbine blade.

Abstract: A turbine shroud has a plurality of shroud segments disposed circumferentially one adjacent to another. Each segment has a flow restrictor projecting integrally from one end face thereof and overlapping a corresponding end face of a circumferentially adjacent segment. The overlap between the circumferentially adjacent segments restricts gas leakage through the inter-segment gap between adjacent shroud segments.

Abstract: A turbine shroud segment is metal injection molded (MIM) about a low melting point material insert. The low melting point material is dissolved using heat during the heat treatment cycle required for the MIM material, thereby leaving internal cooling passages in the MIM shroud segment without extra manufacturing operation.