Abstract: A gas turbine engine comprising a turbine section cooling system and a method of cooling a turbine section of a gas turbine engine is provided. The gas turbine engine comprises in flow series a compressor section, a combustor, and a turbine section, the engine further comprising a turbine section cooling system. The turbine section cooling system including a first compressed air bleed arrangement and a second compressed air bleed arrangement. The first compressed air bleed arrangement bleeds a first flow of compressed air from a high pressure stage of the compressor section. The first flow of compressed air bypasses the combustor and arrives at the turbine section to form a sealing and/or cooling flow at a row of stator vanes upstream of an adjacent rotor disc. The second compressed air bleed arrangement bleeds a second flow of compressed air from one or more lower pressure stages of the compressor section.

Abstract: A gas turbine engine comprising a turbine section cooling system and a method of cooling a turbine section of a gas turbine engine is provided. The gas turbine engine comprises in flow series a compressor section, a combustor, and a turbine section, the engine further comprising a turbine section cooling system. The turbine section cooling system including a first compressed air bleed arrangement and a second compressed air bleed arrangement. The first compressed air bleed arrangement bleeds a first flow of compressed air from a high pressure stage of the compressor section. The first flow of compressed air bypasses the combustor and arrives at the turbine section to form a sealing and/or cooling flow at a row of stator vanes upstream of an adjacent rotor disc. The second compressed air bleed arrangement bleeds a second flow of compressed air from one or more lower pressure stages of the compressor section.

Abstract: An impingement cooling arrangement comprises a projection extending partially across a coolant passage upstream of a jet aperture. An end surface of the projection increases the available surface area for heat exchange with a cross flow whilst a coolant air flow jetted from the jet aperture can transgress a proportion of the air flow passing between the end surface and a junction surface incorporating the jet aperture. A spacing gap B between the end surface and the junction surface avoids localised distortions to the cross flow whilst the projection provides that the coolant air flow projected from the jet aperture mostly passes through a lower turbulence wake downstream of the projection for greater impingement upon a target surface for heat transfer and cooling efficiency. Typically the impingement cooling arrangement is incorporated within turbine blades or vanes of a jet engine.

Abstract: Efficient cooling of a stage of gas turbine engine turbine blades (36) is achieved by first reducing the pressure of the cooling air after it has been bled from the annulus of the compressor (12) by passing it through a diffuser (30), to a pressure magnitude lower than is required at entry to the turbine blades, then re-pressurizing the bled air up to the required entry pressure, by passing it through a radial compressor defined by a cowl (44) positioned in close spaced, co-rotational relationship with the downstream face of the associated turbine disk (34).

Abstract: Rotary apparatus for a gas turbine engine comprises a rotor assembly and first and second stator assemblies mounted coaxially with respect to each other. The first stator assembly is upstream of the second stator assembly, and the second stator assembly is upstream of the rotor assembly. The rotor assembly comprises an annular array of rotor blades, and each stator assembly comprises an annular array of stator vanes. Each vane has a leading edge and a trailing edge. The stator assemblies are circumferentially translatable relative to each other between a first condition and a second condition. In the first condition, at least a part of each vane of the second stator assembly extends beyond the trailing edge of the respective vane of the first stator assembly.

Abstract: Rotary apparatus (17) for a gas turbine engine (10) comprises a rotor assembly (63) and first and second stator assemblies (55, 59) mounted coaxially with respect to each other. Each vane (58, 62) has a leading edge and a trailing edge. The stator assemblies (55, 59) are circumferentially translatable relative to each other between a first condition in which each of the vanes (62) of the second stator assembly (59) is substantially aerodynamically aligned with a respective one of the vanes (58) of the first stator assembly (55), and a second condition in which the vanes (58, 62) of the first and second stator assemblies (55, 59) are out of aerodynamic alignment with each other. In the first condition, at least a part of each vane (62) of the second stator assembly (59) extends beyond the trailing edge of the respective vane (58) of the first stator assembly (55).

Abstract: An aerofoil blade or vane for a gas turbine engine comprises a body member having an inner end for mounting the blade on a shaft and an outer or tip end. A plurality of cooling passages are formed within the blade, the cooling passages comprising a plurality of inlet passages along which cooling air flows from the base towards the tip region of the blade and a plurality of return passages along which cooling air flows from the tip towards the base region of the blade. At least some of the passages are connected by a common chamber located within the tip region of the blade.

Abstract: A gas turbine engine blade or vane comprises inner linked chambers. A chamber adjacent the leading edge is provided with an inlet for receiving cooling fluid and a chamber adjacent the trailing edge is provided with an outlet for exhausting cooling fluid. The chambers are arranged in series from the leading edge to the trailing edge so as to direct cooling fluid within the aerofoil blade or vane from the leading edge region to the trailing edge region.

Abstract: Where gas turbine engine structure eg combustion equipment, is to be air impingement cooled, the surface which receives the air jets is so shaped as to produce boundary layer separation zones 34, 38 and 44 in the cooling air, as it spreads across the surface. Mixing of the boundary layer with the remainder of the air flow results, followed by the re-establishment of the boundary layer. The new boundary layer is cooler than the original layer and so provides more effective cooling.

Abstract: Where gas turbine engine structure eg combustion equipment, is to be air impingement cooled, the surface which receives the air jets is so shaped as to produce boundary layer separation zones 34, 38 and 44 in the cooling air, as it spreads across the surface. Mixing of the boundary layer with the remainder of the air flow results, followed by the re-establishment of the boundary layer. The new boundary layer is cooler than the original layer and so provides more effective cooling.

Abstract: A gas turbine engine blade or vane comprises inner linked chambers. A chamber adjacent the leading edge is provided with an inlet for receiving cooling fluid and a chamber adjacent the trailing edge is provided with an outlet for exhausting cooling fluid. The chambers are arranged in series from the leading edge to the trailing edge so as to direct cooling fluid within the aerofoil blade or vane from the leading edge region to the trailing edge region.

Abstract: An aerofoil blade or vane for a gas turbine engine comprises a body member having an inner end for mounting the blade on a shaft and an outer or tip end. A plurality of cooling passages are formed within the blade, the cooling passages comprising a plurality of inlet passages along which cooling air flows from the base towards the tip region of the blade and a plurality of return passages along which cooling air flows from the tip towards the base region of the blade. At least some of the passages are connected by a common chamber located within the tip region of the blade.

Abstract: Cooling air enters a cooling path 56 from a supply at 58. The path follows a serpentine path consecutively through a number of components to be cooled, including guide vanes 38, 40, blades 30, a shroud 54 and a shaft 18. A valve 68 is provided at the downstream end of the path 56. The setting of the valve 56 modulates the flow and pressure of air along the path 56, without wholly preventing it.

Abstract: A turbine assembly (35) for a gas turbine engine (10) comprises a rotatable support arrangement (38) which comprises means for mounting thereon a plurality of turbine blades (36). The turbine assembly (35) defines flow path means (43) for a flow of cooling fluid therethrough. The flow path means (43) is connectable to a supply of relatively cold cooling fluid. The flow path means 43 is arranged such that the relatively cold cooling fluid is driven radially outwardly through the flow path means (43) substantially wholly by the centrifugal force generated the rotation of the turbine assembly (35) in operation. Relatively hot cooling fluid is displaced by the relatively cold cooling fluid radially inwardly through the flow path means (43).

Abstract: A turbine blade for a gas turbine engine comprises an aerofoil having a suction and pressure side. The pressure side is provided with a reflex curvature at the aerofoil trailing edge region so as to reduce the thickness of the aerofoil in that region.

Abstract: A turbine assembly for a gas turbine engine includes a plurality of turbine blades (32) mounted on a rotatable support means in the form of a turbine disc so as to extend radially therefrom. The turbine blades include circumferentially extending blade platforms (40) spaced from the turbine disc and means are provided for allowing the passage of air between an internal region of the blades (32) and a space located between the blade platforms (40) and the turbine disc. The air may flow out of and back into the same turbine blade, or may flow into an adjacent blade. This flow of air results in the cooling of the blade platforms (40).

Abstract: A turbine blade for a gas turbine engine comprises an aerofoil having a suction and pressure side. The pressure side is provided with a reflex curvature at the aerofoil trailing edge region so as to reduce the thickness of the aerofoil in that region.

Abstract: A gas turbine engine aerofoil (24) has a plurality of attenuation chambers (34) positioned between a cooling air passageway (26) and its leading edge (28). The pressure of the cooling air passing from the passageway (26) to the exterior surface of the leading edge is attenuated by impingement on the opposing walls of the respective chambers (34), or by expansion therein, prior to leaving the chambers(34), via passageways (38).

Abstract: A turbine assembly (35) for a gas turbine engine (10) comprises a rotatable support arrangement (38) which comprises means for mounting thereon a plurality of turbine blades (36). The turbine assembly (35) defines flow path means (43) for a flow of cooling fluid therethrough. The flow path means (43) is connectable to a supply of relatively cold cooling fluid. The flow path means 43 is arranged such that the relatively cold cooling fluid is driven radially outwardly through the flow path means (43) substantially wholly by the centrifugal force generated the rotation of the turbine assembly (35) in operation. Relatively hot cooling fluid is displaced by the relatively cold cooling fluid radially inwardly through the flow path means (43).

Abstract: A turbine assembly for a gas turbine engine includes a plurality of turbine blades (22) mounted on a rotatable support means in the form of a turbine disc so as to extend radially therefrom. The turbine blades include circumferentially extending blade platforms (30) spaced from the turbine disc and means are provided for allowing the passage of air between an internal region of the blades (22) and a space located between the blade platforms (30) and the turbine disc. The air may flow out of and back into the same turbine blade, or may flow into an adjacent blade. This flow of air results in the cooling of the blade platforms (30).