TRAILING EDGE EJECTION COOLING

Abstract

A hollow aerofoil is described having a leading edge and a trailing edge. The leading edge and trailing edge are connected by a pressure surface side (34) and a suction surface side (37) and one or more cavities are bounded by the pressure surface side (34) and/or suction surface side (37). In use, the cavity is arranged to receive coolant from a coolant source. The trailing edge has an apogee (36) where the pressure surface side (34) and suction surface side (37) meet. In an embodiment, a row of holes (32) is provided to a pressure surface side of a centreline of the apogee (36), the holes (32) being in fluid communication with cavity. The arrangement of the holes is such that outlets (33) to the holes extend from the apogee (36) and onto an adjacent part of the pressure surface side (34). A method for the manufacture of the aerofoil is also described.

Claims

1. An aerofoil having a leading edge and a trailing edge, the leading edge and trailing edge connected by a pressure surface side and a suction surface side and one or more cavities bounded by at least one of the pressure surface side or the suction surface side and arranged, in use, to receive coolant from a coolant source, the trailing edge having an apogee where the pressure surface side and the suction surface side meet and wherein a row of holes is provided with the centres of the holes arranged to a pressure surface side or a suction surface side of a centreline of the apogee the holes in fluid communication with one or more of the cavities the arrangement being such that outlets to the holes extend from the apogee and onto an adjacent part of the pressure surface side or the suction surface side.

2. An aerofoil as claimed in claim 1 wherein the row of holes extends along the entire apogee.

3. An aerofoil as claimed in claim 1 wherein the row comprises equally spaced holes.

4. An aerofoil as claimed in claim 1 wherein the row comprises unequally spaced holes.

5. An aerofoil as claimed in claim 1 wherein the holes of the row are grouped with larger spaces between grouped holes.

6. An aerofoil as claimed in claim 1 wherein the holes have a diameter of from 0.3 to 0.6 mm.

7. An aerofoil as claimed in claim 1 wherein the apogee has a thickness of from 0.2 to 0.5 mm.

8. An aerofoil as claimed in claim 1 wherein the cross sectional shape of the holes is selected from; circular, racetrack, elliptical or rectangular.

9. An aerofoil as claimed in claim 1 wherein one or more holes are fanned along a centreline of the hole.

10. A method for manufacturing an aerofoil comprising; manufacturing a blade body having a leading edge and a trailing edge, pressure surface side and a suction surface side and a cavity bounded by at least one of the pressure surface side or the suction surface side and the trailing edge; in an apogee of the trailing edge, drilling holes having their centres arranged to the pressure surface side or the suction surface side of a centreline of the apogee, the holes extending orthogonally to the apogee and into the cavity; after drilling the holes, machine the trailing edge on the pressure surface side or the suction surface side so as to thin the apogee on the side to which the holes are drilled whilst retaining a part of the outlet of the holes in the apogee.

11. A method as claimed in claim 10 wherein the step of drilling the holes involves EDM.

12. A method as claimed in claim 10 wherein the step of drilling involves a method selected from; laser drilling, STEM drilling or manual drilling.

13. A method as claimed in claim 10 wherein the step of machining involves an adaptive machining process.

14. A method as claimed in claim 10 further comprising performing a surface finishing operation to the machined surface.

15. A method as claimed in claim 10 wherein the step of machining the trailing edge results in an apogee thickness of from 0.2 to 0.5 mm.

16. A method as claimed in claim 10, wherein the holes are drilled with a diameter of from 0.3 to 0.6 mm.

Description

BRIEF DESCRIPTION OF DRAWINGS

[0018] Some embodiments of the disclosure will now be further described with reference to the accompanying Figures in which:

[0019] FIG. 1 is a sectional side view of a known gas turbine engine into which aerofoils of the disclosure might usefully be employed;

[0020] FIG. 2 is a sectional view through a trailing edge of a prior known turbine blade design illustrating material removed from the cast blade prior to machining of cooling holes into the trailing edge;

[0021] FIG. 3 is a sectional view through a trailing edge of a turbine blade design in accordance with an embodiment of the disclosure illustrating material removed from the cast blade after machining of cooling holes into the trailing edge;

[0022] FIG. 4 is a schematic showing air flow through and around the trailing edge of FIG. 3;

[0023] FIG. 5 is a sectional view through a trailing edge of a prior known turbine blade after machining;

[0024] FIG. 6 is a sectional view through a trailing edge of a turbine blade in accordance with the disclosure after machining;

[0025] FIG. 7 is a perspective view of a trailing edge bearing one row of cooling holes provided in accordance with the prior art and a second row of cooling holes provided in accordance with an embodiment of the disclosure;

[0026] FIG. 8 shows a cross section of an aerofoil in accordance with an embodiment of the disclosure.

DETAILED DESCRIPTION OF DRAWINGS AND EMBODIMENTS

[0027] With reference to FIG. 1, a gas turbine engine is generally indicated at 10, having a principal and rotational axis 11. The engine 10 comprises, in axial flow series, an air intake 12, a propulsive fan 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, a low-pressure turbine 17 and an exhaust nozzle 18. A nacelle 20 generally surrounds the engine 10 and defines the intake 12.

[0028] The gas turbine engine 10 works in the conventional manner so that air entering the intake 12 is accelerated by the fan 13 to produce two air flows: a first air flow into the high-pressure compressor 14 and a second air flow which passes through a bypass duct 21 to provide propulsive thrust. The high-pressure compressor 14 compresses the air flow directed into it before delivering that air to the combustion equipment 15.

[0029] In the combustion equipment 15 the air flow is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high and low-pressure turbines 16, 17 before being exhausted through the nozzle 18 to provide additional propulsive thrust. The high 16 and low 17 pressure turbines drive respectively the high pressure compressor 14 and the fan 13, each by suitable interconnecting shaft.

[0030] Aerofoils in accordance with the disclosure may usefully be applied in, for example, the turbines 16 and 17 of the described gas turbine engine.

[0031] FIG. 2 shows a trailing edge 1 of a turbine blade known from the prior art. The trailing edge 1 has cooling holes 2 which have an outlet 3 arranged in a pressure surface side 4 of the trailing edge 1. The dotted outline 5 represents the outline of a casting from which the trailing edge 1 is machined. The outlined section 5 is removed from the casting prior to the holes 2 being drilled into the trailing edge from the pressure side surface 4. The trailing edge has an apogee 6 whose thickness is reduced when the outlined section 5 is removed. The apogee 6 is defined by the meeting of the pressure surface side 4 and suction surface side 7.

[0032] FIG. 3 shows a trailing edge 31 of a turbine blade in accordance with an embodiment of the disclosure. The trailing edge 31 has a pressure surface side 34, and a suction surface side 37 which meet to define an apogee 36. A cooling hole 32 has an outlet 33 which extends both in to the apogee 36 and the pressure surface side 34. The outlined section 35 is removed from the casting after the holes 32 have been drilled into trailing edge from the apogee side 36a. As can be seen, in contrast to the prior art shown in FIG. 2, this permits the line of drilling to be perpendicular to the apogee surface 36a. Dotted line 32a represents a wall portion of the hole 32 which is subsequently removed during the machining operation.

[0033] FIG. 4 illustrates air flow in the trailing edge 31 of FIG. 4. The dotted line arrows represent cooling air passing from a cavity within the blade (not shown) and through the trailing edge 31 through cooling hole 32. The solid line arrows represent working fluid flow over the blade. As can be seen cooling air 40 exiting outlet 33 in the apogee 36 fills a void between streams of work fluid passing over the apogee 36 from the pressure surface side 34 and the suction surface side 37. Aerodynamic performance is expected to be improved as aerofoil wake loss is reduced due to this filing of the void.

[0034] FIGS. 5 and 6 show comparable sections of a prior art trailing edge and a trailing edge in accordance with an embodiment of the disclosure. As can be seen, each arrangement has a hole 52, 62 extending through the trailing edge on a pressure surface side 54, 64 of an apogee 56, 66. In the arrangement in accordance with the disclosure, the outlet of the hole 62 extends across part of the apogee 66 and the pressure surface side 64. In the prior art arrangement of FIG. 5, the outlet is present only in the pressure surface side 54.

[0035] FIG. 7 shows a perspective view of the holes of FIGS. 5 and 6 which shows the apogee 76 and pressure surface side 74 of a trailing edge bearing the holes 53, 63. This more clearly illustrates the positioning of the outlets 53 and 63.

[0036] As can be seen, the aerofoil of FIG. 8 has a leading edge 88, and a trailing edge 81. A pressure surface side wall 84 and a suctions surface side wall 87 extend between the leading edge 88 and trailing edge 81. The walls 84, 87 enclose cavities including trailing edge cavity 80. A cooling hole 82 extends from the cavity 80 to an apogee 86 of the trailing edge 81. The outlet 83 to the hole 82 extends into both the pressure surface side wall 84 and the apogee 86.

[0037] When in operation in a gas turbine engine, coolant is delivered into the cavity 80 and directed out through a row of holes 83 extending along the apogee 86 (as shown in the end view of the figure). For example, the coolant may be derived from compressed air which has by-passed the combustor in the engine and is distributed though channels beneath the root of a turbine blade comprising the aerofoil. Rotation in the system generates a radial pressure gradient encouraging flow into the cavities including trailing edge cavity 80 of the aerofoil from which the coolant exits through the holes 82 by outlets 83. On alternative arrangement other sources of coolant may be used and different delivery routes for supplying coolant into the cavity 80 may also be used.