A gas turbine engine is provided that includes compressor and combustor sections, inner and outer casings, an annular diffuser, an inner diffuser casing, a heat exchanger, and an HPT stator vane stage. An annular combustor is disposed radially inward of the outer casing and has inner and outer radial wall structures. The outer casing and the combustor outer radial wall structure define a diffuser OD flow path. The annular diffuser directs diffuser gas towards the combustor section. The inner diffuser casing is disposed radially inward of the annular combustor and spaced apart from the combustor inner radial wall structure. The inner casing is disposed radially inward of and spaced apart from the inner diffuser casing. The inner diffuser casing and the inner casing define an ICF passage. The heat exchanger is configured to produce intercooler gas. Intercooler gas is directed through the ICF passage and into the HPT stator vanes.

A gas turbine engine is provided having an axial centerline, a compressor section, a turbine section, and a combustor section. The turbine section has a turbine first vane assembly that includes an annular first vane inner radial support. The combustor section has an outer casing, an annular combustor, and a unitary inner diffuser structure. The annular combustor has an inner radial flange. The unitary inner diffuser structure includes a compressor discharge, an inner diffuser case, and a tangential onboard injector (TOBI) inseparably attached to one another. The unitary inner diffuser structure further includes an outer radial flange and a TOBI connection flange. The annular FV inner radial support, the combustor inner radial flange, and the TOBI connection flange are secured to one another.

A design for a casing of a large turbine is described with the casing including at least a front section, a middle section and an end section designed such that changes to the mold of the casing required to provide for a change in rotational speed to adapt the turbine to a different power grid frequency are limited to the mold for the middle section of the casing.

The invention relates to a guide vane intended to be mounted in a turbine engine between an inner shroud (17) and an outer shroud (16), comprising a longitudinal straightening body (41) for an air flow extending between a first end (42) intended to be positioned at the inner shroud (17) and a second end (44) intended to be positioned at the outer shroud (16), the longitudinal straightening body having an aerodynamic profile defined by a leading edge (41a) and a trailing edge (41b) in the flow direction of the air flow, and by a camber line (41c) extending from the leading edge (41a) to the trailing edge (41b). It further comprises a first attachment heel (43) and a second attachment heel (45) positioned in the continuation of the longitudinal body (41) at the first end (42) and the second end (44) respectively, the first (43) and second (45) attachment heels being planar elements arranged parallel with respect to one another, each attachment heel (43; 45) being arranged at a distance from the leading edge (41a) and from the trailing edge (41b).

A method for assembling equipment on a support of a turbo-machine. The method comprises the step of mechanically assembling the equipment to a part of the support by means of a fastening device and a further step consisting of providing a coating layer of resin at the interface between the fastening device and the support part. The coating layer allows to perfectly seal the interface between the fastening device and the support part thus avoiding the oxidation of the internal surface without painting.

A method of producing a rotor of a charging apparatus may include the steps of providing at least one compressor wheel and a turbine wheel. The compressor wheel and the turbine wheel may each include a bearing section having a radial bearing surface at a longitudinal end for mounting a bearing housing. At least one of the radial bearing surfaces may include a radial oversizing corresponding to a rotationally asymmetric geometry between at least the bearing section of the compressor wheel and the bearing section of the turbine wheel. The method may include the step of assembling the compressor wheel, the turbine wheel and each bearing section together to form a unitary structure, and machining the at least one of the radial bearing surfaces to reduce the respective radial oversizing until each of the radial bearing surfaces are rotationally symmetrical with respect to each other.

A turbine blade with an integral flow meter is provided. The turbine blade includes a trailing edge and a leading edge opposite the trailing edge. The turbine blade includes at least one cooling passage defined internally within the turbine blade, and the at least one cooling passage is in fluid communication with a source of cooling fluid via an inlet to receive a cooling fluid. The turbine blade also includes at least one flow meter formed within the turbine blade at the inlet that supplies the cooling fluid to the at least one cooling passage.

An extruded profile for manufacturing a blade of an outlet guide vane of a turbine engine. A cross-sectional area has an axial length LAX and a thickness D/LAX relative to the axial length LAX. A cross-sectional area has an at least nearly axisymmetric leading edge region, a first transition region having a varying relative thickness D/LAX. A first constant region has a relative thickness D/LAX at least substantially constant and, relative to a leading edge of the extruded profile, begins at the closest at 10% LAX and ends at the furthest at 50% LAX. A second transition region has a varying relative thickness D/LAX and, relative to the leading edge of the extruded profile, begins at the closest at 30% LAX and ends at the furthest at 90% LAX. A second constant region has a relative thickness D/LAX at least substantially constant and an axial length X of 40% LAX at most; and an at least nearly axisymmetric trailing edge region.

A turbine including a casing and a nozzle including an outer metallic shroud secured to the casing, an inner metallic shroud, and a plurality of nozzle segments made of CMC forming a crown extending between the outer shroud and the inner shroud, each segment including a strut, an inner platform, an outer platform and at least one airfoil having a hollow profile traversed by the strut, wherein for each airfoil, the outer platform includes an axial stop extending in outward radial protrusion from the outer platform, and the outer metallic shroud comprises a complementary axial stop extending in inward radial protrusion from the outer metallic shroud, the axial stop being upstream and axially bearing against the complementary axial stop, and machined with an angle of machining chosen to adjust the orientation of said at least one blade of the segment with respect to the axial direction.

A method for centering an engine structure such as a bearing housing is provided which may be used for example, during assembly of a mid turbine frame or other engine case structure. The method according to one embodiment may include machining spokes with an outer case of the mid turbine frame in situ to eliminate stack-up and then applying the retaining device to retain the spokes with respect to the outer case, thereby assuring the co-axial relationship between the outer case and the bearing housing supported within the mid turbine frame.