Abstract: There is disclosed wall cooling system 50 having a double-wall geometry. A first wall 55 and a second wall 60 extend over a plan area with the second wall spaced from the first wall by a gap. The first wall 55 has multiple upstanding members 65 spanning the gap and contacting the second wall 60 such that the first and second walls are mechanically and thermally connected. The first wall 55 is shaped so as to provide a two-dimensional array of crests 85 and recesses 90. The crests 85 are spaced from the second wall 60. The first wall 55 has a plurality of through-holes 70 for flow of coolant through the first wall and into the gap. The cooling system 50 is suitable for use in a gas turbine engine 10, for example in the turbine 17, 19.

Abstract: A stator assembly comprises an annular platform (1) defining a boundary of an annulus and a plurality of guide vanes (2) for arranging in a circumferential array on the annular platform (1). The platform (1) includes a circumferential array of slots each slot having, at a first end, walls converging (3) from an annulus facing side towards an opposite side of the platform and then extending parallel (4) towards a second end. One or more cooling holes (5) is arranged in a wall (3,4) of the slot. Each vane (2) includes a root portion (7) which converges in a first region (8) distal to the root end and then extends parallel towards the root end in a second region (9). The root portion (7) is configured to engage in a slot of the platform (1) with the first region (8) abutting a convergent wall portion (3) of the slot and the second region (9) spaced from a parallel wall (4) of the slot.

Abstract: A self-sealing impingement cooling tube comprises a perimeter wall (51) defining a first end (54) and a second end (53), for example these may be a trailing edge and leading edge ends of an aerofoil. First and second wall sections extend from the first end (54) to the second end, the first and second wall sections joining at the first end. An adjustable gap (53) resides between the wall sections at the second end. One or both of the first and second wall sections is resiliently deformable such that the first and second wall sections can be positioned closer together and the adjustable gap (53) simultaneously reduced whereby to enable the tube to be inserted into a cavity which has a dimension smaller than the distance between the first and second wall sections when the perimeter wall (51) is not under load.

Abstract: A stator assembly comprises an annular platform (1) defining a boundary of an annulus and a plurality of guide vanes (2) for arranging in a circumferential array on the annular platform (1). The platform (1) includes a circumferential array of slots each slot having, at a first end, walls converging (3) from an annulus facing side towards an opposite side of the platform and then extending parallel (4) towards a second end. One or more cooling holes (5) is arranged in a wall (3,4) of the slot. Each vane (2) includes a root portion (7) which converges in a first region (8) distal to the root end and then extends parallel towards the root end in a second region (9). The root portion (7) is configured to engage in a slot of the platform (1) with the first region (8) abutting a convergent wall portion (3) of the slot and the second region (9) spaced from a parallel wall (4) of the slot.

Abstract: A component of a turbine stage of a gas turbine engine is provided, the component forming an endwall for the working gas annulus of the stage. The component has one or more internal plena behind the endwall which, in use, contain a flow of cooling air. The component further has a plurality of exhaust holes in the endwall. The holes connect the plena to a gas-washed surface of the endwall such that the cooling air effuses through the holes to form a cooling film over the gas-washed surface. Each exhaust hole has a flow cross-sectional area which is greater at an intermediate position between the entrance of the hole from the respective plenum and the exit of the hole to the gas-washed surface than it is at said exit.

Abstract: A shroud segment for a turbine stage of a gas turbine engine forms an endwall for the working gas annulus of the stage. The segment also provides a close clearance to the tips of a row of turbine blades which sweep across the segment. In use a leakage flow of the working gas passes through the clearance gap between the blade tips and the segment. The segment has a plurality of cooling holes and respective air feed passages for the cooling holes. The cooling holes are distributed over that part of the gas-washed surface of the segment which is swept by the blade tips. The cooling holes deliver, in use, cooling air which spreads over the gas-washed surface. The feed passages are configured such that the delivered air opposes the leakage flow of the working gas.

Abstract: An aerofoil (20) for a gas turbine engine (10) is hollow to define an interior volume (24) through which cooling air flows. Passages (26) interconnect the interior volume (24) with the exterior of the aerofoil (20). Each passage (26) is provided with an inlet (32) within the interior volume (24) which is elongated along an axis (36) parallel with the direction of cooling air flow through the interior volume (24). The arrangement reduces any tendency for the passages (24) to block though the build up of dirt particles.

Abstract: A component of a gas turbine engine is provided. The component includes an external wall which, in use, is exposed on one surface thereof to working gas flowing through the engine. The component further includes effusion cooling holes formed in the external wall. In use, cooling air blows through the cooling holes to form a cooling film on the surface of the external wall exposed to the working gas. The component further includes an air inlet arrangement which receives the cooling air for distribution to the cooling holes. The component further includes a plurality of metering feeds and a plurality of supply plena. The metering feeds meter the cooling air from the air inlet arrangement to respective of the supply plena, which in turn supply the metered cooling air to respective portions of the cooling holes.

Abstract: A component of a turbine stage of a gas turbine engine is provided, the component forming an endwall for the working gas annulus of the stage. The component has one or more internal plena behind the endwall which, in use, contain a flow of cooling air. The component further has a plurality of exhaust holes in the endwall. The holes connect the plena to a gas-washed surface of the endwall such that the cooling air effuses through the holes to form a cooling film over the gas-washed surface. Each exhaust hole has a flow cross-sectional area which is greater at an intermediate position between the entrance of the hole from the respective plenum and the exit of the hole to the gas-washed surface than it is at said exit.

Abstract: An aerofoil for a gas turbine engine comprising a pressure wall and a suction wall and defining leading and trailing edges, the walls define a passage into which is supplied a cooling fluid, an array of cooling holes is provided through at least one of the walls to allow the cooling fluid to flow from an interior surface to an exterior surface. The array of holes comprise two groups, the holes of each group are angled to intersect the holes of the other group and are characterized in that the holes of at least one of the groups comprises two or more holes at different angles to one another to vary the porosity of the wall to account for otherwise varying wall temperatures. This arrangement also allows either less coolant mass flow to maintain a constant metal temperature, or a lower metal temperature for a given coolant mass flow.

Abstract: A shroud segment for a turbine stage of a gas turbine engine forms an endwall for the working gas annulus of the stage. The segment also provides a close clearance to the tips of a row of turbine blades which sweep across the segment. In use, a mainstream flow of the working gas passes through the passages formed between adjacent turbine blades. The segment has a plurality of cooling holes and respective air feed passages for the cooling holes. The cooling holes are distributed over that part of the gas-washed surface of the segment which is swept by the blade tips. The cooling holes deliver, in use, cooling air which spreads over the gas-washed surface. The feed passages are configured such that the delivered air has swirl directions which are co-directionally aligned with the swirl directions of the mainstream flow at the segment.

Abstract: A shroud segment for a turbine stage of a gas turbine engine forms an endwall for the working gas annulus of the stage. The segment also provides a close clearance to the tips of a row of turbine blades which sweep across the segment. In use a leakage flow of the working gas passes through the clearance gap between the blade tips and the segment. The segment has a plurality of cooling holes and respective air feed passages for the cooling holes. The cooling holes are distributed over that part of the gas-washed surface of the segment which is swept by the blade tips. The cooling holes deliver, in use, cooling air which spreads over the gas-washed surface. The feed passages are configured such that the delivered air opposes the leakage flow of the working gas.

Abstract: An aerofoil (20) for a gas turbine engine (10) is hollow to define an interior volume (24) through which cooling air flows. Passages (26) interconnect the interior volume (24) with the exterior of the aerofoil (20). Each passage (26) is provided with an inlet (32) within the interior volume (24) which is elongated along an axis (36) parallel with the direction of cooling air flow through the interior volume (24). The arrangement reduces any tendency for the passages (24) to block though the build up of dirt particles.

Abstract: An aerofoil for a gas turbine engine comprising a pressure wall and a suction wall and defining leading and trailing edges, the walls define a passage into which is supplied a cooling fluid, an array of cooling holes is provided through at least one of the walls to allow the cooling fluid to flow from an interior surface to an exterior surface. The array of holes comprise two groups, the holes of each group are angled to intersect the holes of the other group and are characterised in that the holes of at least one of the groups comprises two or more holes at different angles to one another to vary the porosity of the wall to account for otherwise varying wall temperatures. This arrangement also allows either less coolant mass flow to maintain a constant metal temperature, or a lower metal temperature for a given coolant mass flow.

Abstract: An aerofoil 1 includes cooling channels 2, 3, 4, 5 with transfer passages 8 between adjacent channels 2, 3, 4, 5. In normal operation, the constriction in each cooling channel 2, 3, 4, 5 ensures direct through coolant airflow in the direction of arrowheads A. However, when a channel 4 is blocked by a blockage 9 airflow in adjacent channels 2, 5 is forced through the transfer passages 8 in order to provide airflow in the blocked channel 4 and facilitate cooling. This blocked channel airflow is in the direction of arrowheads B. Generally, the cross-section of the transfer passages 8 is determined for conformity with the outlet 7 cross-section of the channel 4 in order to achieve substantial coolant flow balance across the coolant channels 2, 3, 4, 5 of the aerofoil 1.

Abstract: An aerofoil 1 includes cooling channels 2,3,4,5 with transfer passages 8 between adjacent channels 2,3,4,5. In normal operation, the constriction in each cooling channel 2,3,4,5 ensures direct through coolant airflow in the direction of arrowheads A. However, when a channel 4 is blocked by a blockage 9 airflow in adjacent channels 2,5 is forced through the transfer passages 8 in order to provide airflow in the blocked channel 4 and facilitate cooling. This blocked channel airflow is in the direction of arrowheads B. Generally, the cross-section of the transfer passages 8 is determined for conformity with the outlet 7 cross-section of the channel 4 in order to achieve substantial coolant flow balance across the coolant channels 2,3,4,5 of the aerofoil 1.