Device for cooling a wall of a component of a gas turbine

Abstract

A device for cooling a wall of a component of a gas turbine, where a flow passes parallel to the wall, having at least one inflow duct provided in the wall that issues into a recess of the wall for supplying cooling air, wherein a center axis of the inflow duct is aligned to an impact wall of double-convex design inside the recess.

Claims

1. A device for cooling a wall of a component of a gas turbine, where a flow passes parallel to the wall, comprising: a recess positioned in the wall; an impact wall positioned in the recess, the impact wall having a double-convex shape; an inflow duct positioned in the wall that opens into the recess for supplying cooling air into the recess, the inflow duct including a center axis, wherein the center axis of the inflow duct is aligned to the impact wall; wherein the impact wall is shaped as a spheroid dome.

2. The device in accordance with claim 1, wherein the center axis of the inflow duct is arranged parallel or at an angle to a surface of the wall of the component.

3. The device in accordance with claim 1, wherein the impact wall is symmetrical to a symmetry plane including the center axis of the inflow duct and arranged perpendicular to the wall of the component.

4. The device in accordance with claim 1, wherein the center axis of the inflow duct intersects with a center of the impact wall.

5. The device in accordance with claim 1, wherein the wall of the component and the impact wall each further include a thermal barrier coating.

6. The device in accordance with claim 1, wherein a transition area from the impact wall to the wall of the component is free of steps.

7. The device in accordance with claim 1, wherein the recess is shaped as a diffuser.

8. The device in accordance with claim 1, wherein the recess is a singular recess opening toward a surface of the wall of the component.

9. The device in accordance with claim 1, wherein the component is a combustion chamber tile.

10. The device in accordance with claim 1, wherein an angle () is formed between a longer rotational axis of the spheroid dome and the wall, and the angle () is within a range of 5 to 45.

11. The device in accordance with claim 10, wherein an angle () is formed between the center axis and a tangent at an intersection point of the center axis with the impact wall, and the angle () is within a range of 15 to 45.

12. The device in accordance with claim 10, wherein an angle () is formed between the center axis and a tangent at an intersection point of the center axis with the impact wall, and the angle () is 30.

13. The device in accordance with claim 1, wherein an angle () is formed between a longer rotational axis of the spheroid dome and the wall, and the angle () is 20.

14. The device in accordance with claim 13, wherein an angle () is formed between the center axis and a tangent at an intersection point of the center axis with the impact wall, and the angle () is within a range of 15 to 45.

15. The device in accordance with claim 13, wherein an angle () is formed between the center axis and a tangent at an intersection point of the center axis with the impact wall, and the angle () is 30.

16. The device in accordance with claim 1, wherein an angle () is formed between the center axis and a tangent at an intersection point of the center axis with the impact wall, and the angle () is within a range of 15 to 45.

17. The device in accordance with claim 1, wherein an angle () is formed between the center axis and a tangent at an intersection point of the center axis with the impact wall, and the angle () is 30.

Description

(1) The present invention is described in the following on the basis of an exemplary embodiment in light of the accompanying drawing. In the drawing,

(2) FIG. 1 shows a schematic representation of a gas-turbine engine in accordance with the present invention,

(3) FIG. 2 shows a perspective, schematic representation of an exemplary embodiment of the device in accordance with the present invention,

(4) FIG. 3 shows a simplified sectional view, by analogy with FIG. 2,

(5) FIG. 4 shows a simplified top view, by analogy with FIGS. 2 and 3,

(6) FIGS. 5 and 6 show detail sectional views in planes perpendicular to the surface of the wall,

(7) FIG. 7 shows a top view in accordance with the state of the art, and

(8) FIG. 8 shows a top view of the exemplary embodiment in accordance with the present invention.

(9) The gas-turbine engine 10 in accordance with FIG. 1 is a generally represented example of a turbomachine where the invention can be used. The engine 10 is of conventional design and includes in the flow direction, one behind the other, an air inlet 11, a fan 12 rotating inside a casing, an intermediate-pressure compressor 13, a high-pressure compressor 14, a combustion chamber 15, a high-pressure turbine 16, an intermediate-pressure turbine 17 and a low-pressure turbine 18 as well as an exhaust nozzle 19, all of which being arranged about a center engine axis 1.

(10) The intermediate-pressure compressor 13 and the high-pressure compressor 14 each include several stages, of which each has an arrangement extending in the circumferential direction of fixed and stationary guide vanes 20, generally referred to as stator vanes and projecting radially inwards from the core engine casing 21 into an annular flow duct through the compressors 13, 14. The compressors furthermore have an arrangement of compressor rotor blades 22 which project radially outwards from a rotatable drum or disk 26 linked to hubs 27 of the high-pressure turbine 16 or the intermediate-pressure turbine 17, respectively.

(11) The turbine sections 16, 17, 18 have similar stages, including an arrangement of fixed stator vanes 23 projecting radially inwards from the casing 21 into the annular flow duct through the turbines 16, 17, 18, and a subsequent arrangement of turbine rotor blades 24 projecting outwards from a rotatable hub 27. The compressor drum or compressor disk 26 and the blades 22 arranged thereon, as well as the turbine rotor hub 27 and the turbine rotor blades 24 arranged thereon rotate about the engine axis 1 during operation.

(12) FIG. 2 shows in a perspective view part of a wall 25 in accordance with the invention of a component, for example a combustion chamber tile. The wall 25 has a surface along which passes a hot gas flow. To cool the surface of the wall 25, an inflow duct 29 is provided which extends parallel to the surface of the wall 25, as is also shown in FIGS. 3 to 6. The inflow duct 29 has a center axis 31 and is provided in the exemplary embodiment shown with a circular cross-section.

(13) A recess 30 of pocket-like design is provided in the wall 25 and includes a rear wall 33 in which the inflow duct 29 issues. The pocket-like recess 30 furthermore includes side walls 34 arranged at an angle to the center axis 31, such that the recess 30 widens like a diffuser starting from the rear wall 33.

(14) Opposite to the outflow opening of the inflow duct 29, an impact wall 32 is arranged in the recess 30 and is designed in the form of a segment of a spheroid or of the back of a spoon and hence of double-convex design relative to a center axis. The segment of the spheroid is also referred to mathematically as a spheroid dome or cap or hood. FIG. 3 shows a view in a sectional plane including the center axis 31 of the inflow duct 29 and arranged perpendicular to the wall 25 of the component to be cooled. This plane, which forms a symmetry plane for the spheroid 37, is thus the drawing plane in FIG. 3. The reference numeral 36 indicates a rotational axis of the spheroid 37. This is the longer axis of the basic ellipse. The rotational axis 36 is arranged at an angle to the wall 25, which can range from 5 to 45. Preferably, this angle is 20. The spheroid is in the side view in FIG. 3 shown as a dashed line in its remaining part. As FIG. 3 shows, a cooling air jet exiting the outflow duct 29 impacts at an acute angle the impact wall 32 and is evenly deflected to both sides. This angle is formed between the center axis 31 and a tangent 38 at the intersection point of the center axis 31 with the impact wall 32 and is for example between 15 and 45, preferably 30. The cooling air thus passes over the impact wall 32 and then exits straight out of the recess 30. This means that no step or similar is formed. As a result, the flow clings without disruption to the surface of the wall 25.

(15) FIG. 4 once again illustrates the diffuser-like widening of the recess 30.

(16) FIGS. 5 and 6 each show perspective sectional views along sectional planes arranged perpendicular to the surface of the wall 25, as resulting from the assignment of the individual sections from FIGS. 5 and 6. Here, the double-convex design of the impact wall can again be discerned in particular. It can furthermore be discerned that the impact wall 32 is symmetrical to a symmetry plane including the center axis 31 and arranged perpendicular to the surface of the wall 25.

(17) FIGS. 7 and 8 each show arrangements of recesses 30 provided in a wall 25 of a component. FIG. 7 shows here an embodiment according to the state of the art, in which the basic surfaces 35 of the recesses 30 are designed smooth or flat, while FIG. 8 shows an embodiment in accordance with the invention with double-convex impact walls 32. From FIGS. 7 and 8 it can be discerned that the individual recesses 30 are not connected to one another, but are placed in a regular arrangement relative to one another in order to achieve an even formation of a cooling film on the surface of the wall 25.

(18) The exemplary embodiments shown dispense, for the purposes of greater clarity, with the illustration of a thermal barrier coating that can be applied on the surface of the wall 25 and at least on the basic surface 35 and the impact wall 32 as a spray coating.

LIST OF REFERENCE NUMERALS

(19) 1 Engine axis 10 Gas-turbine engine/core engine 11 Air inlet 12 Fan 13 Intermediate-pressure compressor (compressor) 14 High-pressure compressor 15 Combustion chamber 16 High-pressure turbine 17 Intermediate-pressure turbine 18 Low-pressure turbine 19 Exhaust nozzle 20 Guide vanes 21 Core engine casing 22 Compressor rotor blades 23 Stator vanes 24 Turbine rotor blades 25 Wall 26 Compressor drum or disk 27 Turbine rotor hub 28 Exhaust cone 29 Inflow duct 30 Recess 31 Center axis 32 Impact wall 33 Rear wall 34 Side wall 35 Basic surface 36 Rotational axis 37 Spheroid 38 Tangent