Abstract: An air-debris separator includes a center body, an inner wall, an outer wall, a separation passage, a first outlet passage and a second outlet passage. The center body extends longitudinally along a longitudinal centerline. The inner wall and the outer wall extend longitudinally along and circumferentially about the center body. The separation passage extends longitudinally within the air-debris separator to the first outlet passage and the second outlet passage. The separation passage is radially between the center body and the outer wall. The separation passage includes a convergent section and a radial outer diameter that decreases as the convergent section extends longitudinally towards the first outlet passage and the second outlet passage. The first outlet passage is radially between the center body and the inner wall. The second outlet passage is formed by and radially between the inner wall and the outer wall.

Abstract: A gas turbine engine is provided that includes a high pressure compressor (HPC), a combustor section, a high pressure turbine (HPT), and a bypass tangential on board injector (TOBI) system. The combustor section has a combustor. A core gas path extends through the HPC, the combustor section, and the HPT. The bypass TOBI system extends circumferentially around the engine axial centerline, and has a plurality of nozzles, inner and outer radial sides, a plurality of first type and second type radial passages configured to allow the gas from the HPC to pass from the inner radial side of the bypass TOBI system to the outer radial side of the bypass TOBI system, wherein the first type radial passages are differently configured from the second type radial passages.

Abstract: A gas turbine engine includes a compressor, a turbine, and a rotational assembly. The compressor and the turbine form a core flow path. The rotational assembly includes a shaft, a compressor rotor, and a turbine rotor. The shaft interconnects the compressor rotor and the turbine rotor. The compressor rotor includes bladed rotor disks. The bladed rotor disks are spaced from the shaft to form an unobstructed passage. The bladed rotor disks form at least one forward rotor cavity and at least one aft cavity. The at least one forward rotor cavity and the at least one aft rotor cavity are connected in fluid communication by the unobstructed passage. The rotational assembly forms a thermal conditioning assembly. The thermal conditioning assembly includes air apertures formed by at least one of the bladed rotor disks. The air apertures extend from the core flow path to the at least one aft rotor cavity.

Abstract: A gas turbine engine is provided that includes a high pressure compressor (HPC), a combustor section, a high pressure turbine (HPT), and a bypass tangential on board injector (TOBI) system. The combustor section has a combustor. A core gas path extends through the HPC, the combustor section, and the HPT. The bypass TOBI system extends circumferentially around the engine axial centerline, and has a plurality of nozzles, inner and outer radial sides, a plurality of first type and second type radial passages configured to allow the gas from the HPC to pass from the inner radial side of the bypass TOBI system to the outer radial side of the bypass TOBI system, wherein the first type radial passages are differently configured from the second type radial passages.

Abstract: A gas turbine engine is provided that includes a high pressure compressor (HPC), a combustor section, a high pressure turbine (HPT), and a bypass tangential on board injector (TOBI) system. The combustor section has a combustor. A core gas path extends through the HPC, the combustor section, and the HPT. The bypass TOBI system extends circumferentially around the engine axial centerline, and has a plurality of nozzles, inner and outer radial sides, a plurality of first type and second type radial passages configured to allow the gas from the HPC to pass from the inner radial side of the bypass TOBI system to the outer radial side of the bypass TOBI system, wherein the first type radial passages are differently configured from the second type radial passages.

Abstract: An aspect includes a system including a high compressor of a gas turbine engine having a ratio of a cold-rotor build clearance to a span between 0.7% and 7%. The cold-rotor build clearance is defined for a plurality of rotor blades of the high compressor with respect to an engine casing assembly interior surface of the high compressor, and the span is defined as a gap between a rotor disk of the high compressor and the engine casing assembly interior surface of the high compressor for at least a last two stages of the high compressor closest to a combustor section of the gas turbine engine. The system also includes at least two bowed rotor management systems for the gas turbine engine to prevent damage to the rotor blades for a bowed rotor condition of the high compressor under a plurality of operating conditions.

Abstract: An aspect includes a system including a high compressor of a gas turbine engine having a ratio of a cold-rotor build clearance to a span between 0.7% and 7%. The cold-rotor build clearance is defined for a plurality of rotor blades of the high compressor with respect to an engine casing assembly interior surface of the high compressor, and the span is defined as a gap between a rotor disk of the high compressor and the engine casing assembly interior surface of the high compressor for at least a last two stages of the high compressor closest to a combustor section of the gas turbine engine. The system also includes at least two bowed rotor management systems for the gas turbine engine to prevent damage to the rotor blades for a bowed rotor condition of the high compressor under a plurality of operating conditions.

Abstract: A splitter for protecting parts of a gas turbine engine is provided. The splitter segregates the flow of hot and cold gases into the air bleed manifold around the high compressor section to reduce maximum temperatures and minimize the temperature gradients of engine parts, especially inner case flanges The splitter divides the air bleed manifold into an annular inner cavity and an outer cavity. The splitter directs relatively cold air from the high compressor bleed ducts into the inner cavity and relatively hot air from the aft hub into the outer cavity, away from the inner case flanges.

Abstract: A disclosed gas turbine engine includes a case for a case for a compressor section including a bleed air slot. The bleed air slot includes an inlet having a first area radially inward of an outlet having a second area with the second area being greater than the first area. The bleed air slot further includes a center portion disposed along a radial line extending from an axis of the engine and an elongated portion extending from the opening at an angle relative to a line normal to the radial line.

Abstract: A splitter for protecting parts of a gas turbine engine is provided. The splitter segregates the flow of hot and cold gases into the air bleed manifold around the high compressor section to reduce maximum temperatures and minimize the temperature gradients of engine parts, especially inner case flanges The splitter divides the air bleed manifold into an annular inner cavity and an outer cavity. The splitter directs relatively cold air from the high compressor bleed ducts into the inner cavity and relatively hot air from the aft hub into the outer cavity, away from the inner case flanges.

Abstract: A disclosed gas turbine engine includes a case for a case for a compressor section including a bleed air slot. The bleed air slot includes an inlet having a first area radially inward of an outlet having a second area with the second area being greater than the first area. The bleed air slot further includes a center portion disposed along a radial line extending from an axis of the engine and an elongated portion extending from the opening at an angle relative to a line normal to the radial line.