EP2505787A1

EP2505787A1 - Component of a gas turbine engine and corresponding gas turbine engine

Info

Publication number: EP2505787A1
Application number: EP12157214A
Authority: EP
Inventors: Lynne Turner; Peter Ireland; Ian Tibbott; Anthony Rawlinson
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2011-03-28
Filing date: 2012-02-28
Publication date: 2012-10-03
Also published as: GB201105105D0; US20120251295A1

Abstract

A component of a gas turbine engine is provided. The component includes an external wall (40) which, in use, is exposed on one surface (41) thereof to working gas flowing through the engine. The component further includes effusion cooling holes formed in the external wall (42). 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 (41). The component further includes an air inlet arrangement (43) which receives the cooling air for distribution to the cooling holes (42). The component further includes a plurality of metering feeds (47) and a plurality of supply plena (45). The metering feeds (47) meter the cooling air from the air inlet arrangement (43) to respective of the supply plena (45), which in turn supply the metered cooling air to respective portions of the cooling holes (42).

Description

The present invention relates to a cooled component for use in gas turbine engines.
With reference to Figure 1, a ducted fan gas turbine engine generally indicated at 10 has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, and intermediate-pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23.
The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 14 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low- pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.
The performance of gas turbine engines, whether measured in terms of efficiency or specific output, is improved by increasing the turbine gas temperature. It is therefore desirable to operate the turbines at the highest possible temperatures. For any engine cycle compression ratio or bypass ratio, increasing the turbine entry gas temperature produces more specific thrust (e.g. engine thrust per unit of air mass flow). However as turbine entry temperatures increase, the life of an un-cooled turbine falls, necessitating the development of better materials and the introduction of internal air cooling.
In modern engines, the high-pressure turbine gas temperatures are hotter than the melting point of the material of the blades and vanes, necessitating internal air cooling of these airfoil components. During its passage through the engine, the mean temperature of the gas stream decreases as power is extracted. Therefore, the need to cool the static and rotary parts of the engine structure decreases as the gas moves from the high-pressure stage(s), through the intermediate-pressure and low-pressure stages, and towards the exit nozzle.
Figure 2 shows an isometric view of a typical single stage cooled turbine. Cooling air flows are indicated by arrows.
Internal convection and external films are the prime methods of cooling the gas path components - airfoils, platforms, shrouds and shroud segments etc. High-pressure turbine nozzle guide vanes 31 (NGVs) consume the greatest amount of cooling air on high temperature engines. High-pressure blades 32 typically use about half of the NGV flow. The intermediate-pressure and low-pressure stages downstream of the HP turbine use progressively less cooling air.
The high-pressure turbine airfoils are cooled by using high pressure air from the compressor that has by-passed the combustor and is therefore relatively cool compared to the gas temperature. Typical cooling air temperatures are between 800 and 1000 K, while gas temperatures can be in excess of 2100 K.
The cooling air from the compressor that is used to cool the hot turbine components is not used fully to extract work from the turbine. Therefore, as extracting coolant flow has an adverse effect on the engine operating efficiency, it is important to use the cooling air effectively.
Ever increasing gas temperature levels combined with a drive towards flatter combustion radial profiles, in the interests of reduced combustor emissions, have resulted in an increase in local gas temperature experienced by the extremities of the blades and vanes, and the working gas annulus endwalls formed e.g. by the NGV 31 inner and outer platforms 33, the blade 32 platform 34, and the blade shroud 35. Effusion cooling holes 36 can be formed in the endwalls of such components, so that cooling air can blow therethrough to form a cooling film over the surface of the endwall exposed to the working gas.
The pressure field into which the cooling film is introduced typically decreases from the leading edge to the trailing edge of the endwall of the component. Thus when the film is supplied from a single pressure source, the film blowing rate and effectiveness at the trailing edge is determined by a need to provide a safe pressure margin at the leading edge. This leads to higher blowing rates at the trailing edge than required, which compromises the balance between film effectiveness and system mass flow rate.
Accordingly, a first aspect of the present invention provides a component of a gas turbine engine, the component including:

an external wall which, in use, is exposed on one surface thereof to working gas flowing through the engine,
effusion cooling holes formed in the external wall, in use cooling air blowing through the cooling holes to form a cooling film on the surface of the external wall exposed to the working gas, and
an air inlet arrangement which receives the cooling air for distribution to the cooling holes;

a plurality of metering feeds and a plurality of supply plena, the metering feeds metering the cooling air from the air inlet arrangement to respective of the supply plena, and the supply plena supplying the metered cooling air to respective portions of the cooling holes.

Advantageously, the metering feeds can be configured to provide different cooling air pressures in the supply plena. In this way, the supply plena and their metering feeds allow the cooling air blown through the cooling holes to be driven by different source pressures. The film blowing rate at different positions on the external wall can thus be matched to the working gas pressure field, leading to enhanced film effectiveness and reduced aerodynamic losses.
The component may have any one or, to the extent that they are compatible, any combination of the following optional features.
Different flow cross-sectional areas of the metering feeds can provide the different cooling air pressures in the supply plena.
The supply plena may be partially defined by the other surface of the external wall. The metering feeds can then be configured to form impingements jets from the cooling air metered therethrough, the impingements jets impinging on said other surface. The jets can provide further cooling of the external wall.
Alternatively, the component may further include a plurality of secondary plena which are partially defined by the other surface of the external wall, and are each in flow series between a respective supply plenum and its respective portion of the cooling holes. In such an arrangement, the supply plena can still provide different source pressures for the cooling holes. However, the component may also include a plurality of jet-forming passages which deliver the cooling air from the supply plena to the secondary plena, the jet-forming passages being configured to form impingements jets from the cooling air delivered therethrough, and the impingements jets impinging on said other surface. Thus the metering feeds do not then have to perform a jet-forming function.
The component may further include an inlet plenum in flow series between the air inlet arrangement and the supply plena, the metering feeds feeding the cooling air from the inlet plenum to respective of the supply plena. The inlet plenum can help to ensure an even distribution of cooling air into the supply plena. However, a different option is to arrange for the entrances to the metering feeds to form directly the cooling air inlet arrangement, e.g. by extending the metering feeds to the rear of the segment.
Typically, in use, the pressure of the working gas to which the external wall is exposed varies from a leading edge to a trailing edge of the external wall, the supply plena, metering feeds and cooling holes being configured such that the pressure of the cooling air blowing through each cooling hole matches the local pressure of the working gas at that cooling hole. For example, the working gas can vary from a higher pressure to a lower pressure from the leading edge to the trailing edge of the external wall.
The cooling holes may be angled in the external wall to further reduce aerodynamic losses. The holes may have fan-shaped or conical exit geometries, e.g. to improve spreading of the cooling film and to reduce the exit velocity of the blown cooling air.
The cooling holes also effect cooling of the external wall by heat transfer from the walls of the holes to the air blowing therethrough. Increasing the cooling hole internal surface roughness can thus enhance cooling effectiveness, as can increasing the lengths of the cooling holes (e.g. by angling the holes and/or increasing the external wall thickness).
The component may provide an endwall to the working gas annulus of the engine, the external wall being the endwall. For example, the component can be a shroud segment, and in particular a high-pressure or intermediate-pressure shroud segment. Alternatively, the component can be a turbine blade or a vane, a platform of the blade or vane forming the endwall, and in particular a high-pressure turbine blade, or a high-pressure or intermediate-pressure nozzle guide vane.
Alternatively, the component can be a combustor, the external wall at least partially defining a combustion chamber of the combustor.
A second aspect of the present invention provides gas turbine engine having one or more components according to the previous aspect.
Embodiments of the invention will now be described by way of example with reference to the accompanying drawings in which:

Figure 1 shows a schematic longitudinal cross-section through a ducted fan gas turbine engine;
Figure 2 shows an isometric view of a typical single stage cooled turbine;
Figure 3 shows a schematic longitudinal cross-section through a first embodiment of a shroud segment for a high-pressure or intermediate-pressure turbine stage of a gas turbine engine; and
Figure 4 shows (a) a second embodiment of a schematic longitudinal cross-section through another shroud segment, and, and (b) a plot of axial variation of the gas path static pressure adjacent the endwall of the segment of the first or second embodiment.

Figure 3 shows a first embodiment of a schematic longitudinal cross-section through a shroud segment for a high-pressure or intermediate-pressure turbine stage of a gas turbine engine. The segment provides an endwall 40 to the working gas annulus with an external gas washed surface 41 that is exposed to the working gas flowing through the engine. Cooling air is blown through a plurality of effusion cooling holes 42 formed in the endwall to form a cooling film over the gas washed surface that protects the endwall from the working gas. Heat transfer from the walls of the holes to the air blowing therethrough also cools the endwall. To prevent working gas being ingested into the segment through the holes, the source pressure of the cooling air must exceed that of working gas.
The cooling air is typically compressed air bled from the compressor section of the engine and bypassing the combustor. The direction of flow of the cooling air is indicated by arrows in Figure 3. The air enters the segment through an air inlet aperture 43 at the rear of the segment and fills an inlet plenum 44. A plurality of supply plena 45 are arranged between the inlet plenum and the endwall 40, with the inner surface 46 of the endwall partially defining the supply plena, and the individual supply plena being separated from each other by internal walls 50. The cooling air is fed into the supply plena from the inlet plenum via respective metering feeds 47. The inlet plenum evens the distribution of the air flow to the supply plena, the pressure (P1-P8) of the cooling air in each supply plenum being a function of at least the flow cross-sectional area of the respective feed into that plenum.
Each supply plenum 45 then supplies the cooling air for a respective portion of the cooling holes 42. In particular, as shown in Figure 3, the supply plena are arranged in a line from the leading edge 48 to the trailing edge 49 of the endwall 40, with each supply plena supplying a respective row of cooling holes. The row by row hole diameter and the number of holes in each row can vary depending on the cooling duty. Each supply plena operates at a different pressure as determined by its metering feed 47. This allows local target pressure margins to be maintained above the working gas path pressure seen by the gas washed surface 41. In particular, axial variation of the gas path static pressure distribution can be accommodated to provide a more uniform blowing rate and mass flow into the cooling film.
Optionally, the metering feeds 47 can form impingement jets from the metered cooling air. These jets then impinge on the inner surface 46 of the endwall to enhance heat transfer from the endwall.
However, another option is to separate the metering and jet forming functionality. Accordingly, Figure 4(a) shows a second embodiment of a schematic longitudinal cross-section through another shroud segment. Corresponding features have the same reference numbers in Figures 3 and 4(a). Figure 4(b) shows a plot of axial variation of the gas path static pressure adjacent the endwall 40 of the segment of the first or second embodiment. The position of the tips of the turbine blades which sweep across the endwall is indicated.
In the second embodiment, each supply plenum 45 is fed by a respective metering feed 47, as in the first embodiment. The positions of the metering feeds are only schematically indicated in Figure 4(a). However, the entrances of the feeds, which are at the rear side of the segment, directly form an air inlet arrangement into the segment, i.e. a separate air inlet aperture and intermediate inlet plenum are not needed. Further, a row of secondary plena 51 are provided between the supply plena and the endwall 40, with the inner surface 46 of the endwall now partially defining the secondary plena. The secondary plena, like the supply plena, are separated from each other by internal walls 52. Each supply plenum has a respective secondary plenum, with jet-forming passages 53 delivering the cooling air from the supply plena to the secondary plena. The jets of cooling air produced by these passages impinge on the inner surface 46 of the endwall to enhance heat transfer from the endwall. Thus the metering feeds 47 ultimately determine the cooling air pressure in the secondary plena, but do not form the impingement jets.
In both embodiments, the provision of supply plena 45 and metering feeds 47 allows better control of mass flow rate through individual rows of cooling holes 42. In this way, a shroud segment can achieve a high level of cooling film effectiveness, as the film can be introduced onto the gas washed surface 41 with a momentum which matches the pressure of the gas in contact with the wall. By matching the cooling air blowing rate to the local gas pressure, engine specific fuel consumption can be reduced as less cooling flow is required. Further, improved cooling film effectiveness allows higher turbine entry temperatures to be achieved.
Although described above in relation to shroud segments, the present invention may also be applied to e.g. a platform of a high-pressure turbine blade, or the platforms of a high-pressure or intermediate-pressure nozzle guide vane. The present invention may additionally be applied to combustor chamber wall cooling.
While the invention has been described in conjunction with the exemplary embodiments described above, many equivalent modifications and variations will be apparent to those skilled in the art when given this disclosure. Accordingly, the exemplary embodiments of the invention set forth above are considered to be illustrative and not limiting. Various changes to the described embodiments may be made without departing from the spirit and scope of the invention.

Claims

A component of a gas turbine engine, the component including:
an external wall (40) which, in use, is exposed on one surface (41) thereof to working gas flowing through the engine,

effusion cooling holes (42) formed in the external wall, in use, cooling air blowing through the cooling holes to form a cooling film on the surface of the external wall exposed to the working gas, and

an air inlet arrangement (43) which receives the cooling air for distribution to the cooling holes;
wherein the component further includes:
a plurality of metering feeds (47) and a plurality of supply plena (45), the metering feeds metering the cooling air from the air inlet arrangement to respective of the supply plena, and the supply plena supplying the metered cooling air to respective portions of the cooling holes.
A component according to claim 1, wherein the supply plena are partially defined by the other surface (46) of the external wall.
A component according to claim 2, wherein the metering feeds are configured to form impingements jets from the cooling air metered therethrough, the impingements jets impinging on said other surface.
A component according to claim 1, further including a plurality of secondary plena (51) which are partially defined by the other surface of the external wall, and are each in flow series between a respective supply plenum and its respective portion of the cooling holes.
A component according to claim 4, further including a plurality of jet-forming passages (53) which deliver the cooling air from the supply plena to the secondary plena, the jet-forming passages being configured to form impingements jets from the cooling air delivered therethrough, and the impingements jets impinging on said other surface.
A component according to any one of the previous claims which further includes an inlet plenum (44) in flow series between the air inlet arrangement and the supply plena, the metering feeds feeding the cooling air from the inlet plenum to respective of the supply plena.
A component according to any one of the previous claims, wherein, in use, the pressure of the working gas to which the external wall is exposed varies from a leading edge (48) to a trailing edge (49) of the external wall, the supply plena, metering feeds and cooling holes being configured such that the pressure of the cooling air blowing through each cooling hole matches the local pressure of the working gas at that cooling hole.
A component according to claim 7, wherein the working gas varies from a higher pressure to a lower pressure from the leading edge to the trailing edge of the external wall.
A component according to any one of the previous claims which provides an endwall to the working gas annulus of the engine, the external wall being the endwall.
A component according to claim 9 which is a shroud segment.
A component according to claim 9 which is a turbine blade or a vane, a platform of the blade or vane forming the endwall.
A component according to any one of claims 1 to 9 which is a combustor, the external wall at least partially defining a combustion chamber of the combustor.
A gas turbine engine having one or more components according to any one of the previous claims.