A compressor rotor of a gas turbine engine includes a rotor body having a face adapted to face an adjacent rotor. The rotor body extends radially between an outer peripheral rim surface and an inner rim surface. The inner rim surface defines a bore of the rotor body. A plurality of blades extends radially from the outer peripheral rim surface. A plurality of anti-vortex fins extends axially from the face of the rotor body facing the adjacent rotor. The plurality of anti-vortex fins forms a plurality of open radial passageways. The plurality of anti-vortex fins extends axially to a predetermined thickness such that, when assembled with the second rotor, axial extremities of the plurality of anti-vortex fins being in close proximity with the adjacent rotor and the adjacent rotor closes the radial passageways. A method of providing a first rotor for assembly with a second facing rotor of a compressor rotor assembly is also presented.

The invention relates to an expansion machine (20), comprising an output shaft (24) and a shaft sealing ring (25) that interacts with the output shaft. The expansion machine (20) has an inflow region (21) and an outflow region (22). During operation, a working medium flows through the expansion machine (20), wherein compressed working medium flows into the inflow region (21) and expanded working medium flows out of the outflow region (22). The shaft sealing ring (25) separates a valve space (11) filled with working medium from a surrounding space (40). A valve (10) is arranged in the expansion machine (20). The pressure in the valve space (11) can be controlled by means of the valve (10).

A turbocharger (10) includes a shaft (26) rotatably supported within a bearing housing (28), a turbine wheel (22) connected to the shaft (26), and a heat shield (150, 250) disposed between the turbine wheel (22) and the bearing housing (28). The heat shield (150, 250) includes surface features (80, 180) formed on at least one of a sidewall (57) portion thereof and a flange (63) portion thereof that locate the heat shield (150, 250) relative to the bearing housing (28) such that the heat shield (150, 250) is coaxial with the rotational axis of the shaft (26).

An aspirating seal assembly for use in a turbine engine is provided. The aspirating seal assembly includes a face seal and a rotary component. The face seal includes a first annular seal surface, and the rotary component includes a second annular seal surface positioned adjacent the first annular seal surface and defining a seal interface therebetween. The face seal is configured to discharge a flow of air towards the seal interface. The seal assembly also includes a first seal member extending between the first and second annular seal surfaces such that the flow of air induces a back pressure across the seal interface, and a second seal member positioned radially inward from the first seal member and extending between the first and second annular seal surfaces. A length of the second seal member is selected to increase the back pressure induced across the seal interface.

A rotor disc for a gas turbine engine includes an annular disc body configured to support a circumferential array of blades and having a plurality of passages defined therethrough. The passages form coils within the disc body and/or have a packing density of at least 0.1 in cross-sectional plane containing the central axis, the packing density being defined by a ratio between an open area of the passages and a solid area of the disc in the cross-sectional plane. A method of manufacturing a rotor disc is also discussed.

This invention concerns a system for cooling components in a gas turbine engine, the gas turbine engine including a compressor for driving a primary gas flow to a combustor and a turbine arranged to be driven by combustion gases from the combustor, wherein the system includes: an annular cooling flow passage arranged for fluid communication between the compressor and the turbine, the flow passage having a first inlet arranged to receive gas from the primary gas flow downstream of compressor, and a second inlet located upstream of the first inlet, wherein the annular cooling flow passage has at least one internal wall for guiding airflow from the first inlet towards the airflow from the second inlet, the airflow from the first and second inlets coalesce within the annular flow passage prior to passing along the passage in a direction from the compressor to the turbine.

A compressor rotor of a gas turbine engine includes a rotor body having a face adapted to face an adjacent rotor. The rotor body extends radially between an outer peripheral rim surface and an inner rim surface. The inner rim surface defines a bore of the rotor body. A plurality of blades extends radially from the outer peripheral rim surface. A plurality of anti-vortex fins extends axially from the face of the rotor body facing the adjacent rotor. The plurality of anti-vortex fins forms a plurality of open radial passageways. The plurality of anti-vortex fins extends axially to a predetermined thickness such that, when assembled with the second rotor, axial extremities of the plurality of anti-vortex fins being in close proximity with the adjacent rotor and the adjacent rotor closes the radial passageways. A method of providing a first rotor for assembly with a second facing rotor of a compressor rotor assembly is also presented.

A solution heat treatment method is disclosed for manufacturing metallic components of a turbo machine, which components provide a hot gas flow channel when assembled in the turbo machine after manufacturing, wherein the components are subjected to a time-temperature-cycle in a furnace. The method includes positioning the components in the furnace in a same principle as the component assembly in the turbo machine, but leaving flow areas and gaps between neighbouring components; then starting the time-temperature-cycle; and applying an inert gas during the solution heat treatment process so that the inert gas flows through flow areas and gaps for achieving a uniform temperature.

There is provided a seal assembly comprising: a first component and a second component spaced apart from the first component so as to define a passage for the transfer of fluid from an inlet of the seal assembly to an outlet of the seal assembly, wherein the first component comprises a concavity at least partially defining the passage, and wherein no part of the second component extends into the portion of the passage bounded by the concavity.

A compressor rotor of a gas turbine engine includes a rotor body having a face adapted to face an adjacent rotor. The rotor body extends radially between an outer peripheral rim surface and an inner rim surface. The inner rim surface defines a bore of the rotor body. A plurality of blades extends radially from the outer peripheral rim surface. A plurality of anti-vortex fins extends axially from the face of the rotor body facing the adjacent rotor. The plurality of anti-vortex fins forms a plurality of open radial passageways. The plurality of anti-vortex fins extends axially to a predetermined thickness such that, when assembled with the second rotor, axial extremities of the plurality of anti-vortex fins being in close proximity with the adjacent rotor and the adjacent rotor closes the radial passageways. A method of providing a first rotor for assembly with a second facing rotor of a compressor rotor assembly is also presented.