GB2523140A - Gas turbine engine component - Google Patents

Abstract

A gas turbine engine component such as a high pressure turbine nozzle guide vane (NGV) or blade includes an external wall which is exposed to hot gas flowing through the engine. The component includes an air inlet arrangement 42 and effusion (film) cooling holes 46 formed in the external wall. Air is supplied from the inlet arrangement 42 via metering / impingement holes 45 to supply plena 43, 44. The plena 43, 44 supply air to the film cooling holes 46. Cooling air flowing through metering holes 45 impinges on the inner surface of the outer wall to enhance heat transfer. Each plenum 43, 44 is supplied by only one or two metering holes, and supplies only one or two effusion cooling holes 46. This allows flow rates to be accurately matched to cooling requirements at each point on the blade surface.

Description

GAS TURBINE ENGINE COMPONENT

The present invention relates to a cooled component for use in gas turbine engines.

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 aerofoil 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 1 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 -aerofoils, 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 aerofoils 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 and overall pressure ratios combined with a drive towards flatter combustion radial profiles have resulted in an increase in local gas temperature experienced by 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.

The aerofoils themselves have cooling schemes that are designed to fit within the aerofoil shape, and to provide cooling both internally (augmented channel flow and/or impingement) and externally (film cooling). Up until recently, the internal cooling schemes featured an internal passage or plenum that received coolant from a central cavity, and then fed the coolant to both the suction and pressure surfaces of the aerofoil. As a result, the internal heat transfer and film cooling blowing rates were not optimised for either surface.

More recent cooling scheme developments have seen the internal plenum divided into two in order to better match the required feed conditions to the pressure or suction sides of the aerofoil. These changes have proved beneficial, whether the internal heat transfer has been achieved by augmented channel flow or by the use of impingement jet cooling.

Nonetheless, each plenum has only one total pressure maintained at a value somewhere between a central cavity pressure of the aerofoil and the external static pressure in the vicinity of the plenum. As each plenum can be relatively long and have many impingement holes (or other type of feed) linking it to the central aerofoil cavity, and as there may be many film cooling holes connecting each plenum to the external surface of the aerofoil, this intermediate plenum pressure can in general only be optimised for one of its film cooling hole and one of its impingement jet. It follows that some of the impingement jets tend to produce higher levels of heat transfer than required and some lower levels. Similarly, some of the film cooling holes deliver larger quantities of flow at high blowing rates than required while others under flow.

Plena embedded into aerofoil or platform walls where the curvature changes rapidly, such as in the vicinity of the aerofoil leading edge, can cause problems in manufacture. The ceramic cores used in the casting of the component tend to crack at the features (termed pedestals) that link the plena to the central core. The pedestals are comparatively fragile and can break in the core injection tool. They can also crack when the molten metal is poured in the casting mould. The local curvature of the aerofoil adds to the problem because the external and internal walls cool down and solidify at different rates -locking in stresses in the pedestals causing them to crack more readily.

Accordingly, in a first aspect the present invention provides a component of a gas turbine engine, the component including: an external wall which, in use, is exposed on the outer 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 an air inlet arrangement which receives the cooling air for distribution to the cooling holes; wherein the component further includes a plurality of metering feeds and a plurality of supply plena, the metering feeds metering the cooling air from the air inlet arrangement to the supply plena, and the supply plena supplying the metered cooling air to the cooling holes; and wherein each supply plenum is fed by just one or two of the metering feeds, and supplies the metered cooling air to just one or two of the cooling holes.

In this way, the supply plena can be kept to a small size, allowing them to be individually tailored to the convection and film cooling conditions associated with the local heat load. In particular, small plena enable the heat transfer engineer to optimise convective heat transfer rates, while simultaneously allowing the film cooling blowing rate to be optimised to provide maximum film coverage and hence film effectiveness. Better control over the film cooling blowing rate and film flow levels can reduce the aerodynamic losses associated with mixing the cooling flow with the mainstream gas flow, and ultimately improve engine efficiency and specific fuel consumption.

Small supply plena can also improve casting yields by reducing core breakage associated with pedestals that form the metering feeds.

Small supply plena can also help to reduce the thermal stress in the walls of the component, particularly near the leading edge of an aerofoil. In this way, the thermal fatigue life of the component can be improved In a second aspect, the present invention provides a gas turbine engine having one or more components according to the first aspect.

Optional features of the invention will now be set out. These are applicable singly or in any combination with any aspect of the invention.

The supply plena may be partially defined by the inner 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 the inner surface.

The component may further include an internal cavity in flow series between the air inlet arrangement and the supply plena, the metering feeds feeding the cooling air from the internal cavity to the supply plena.

The supply plena may be arranged in a regular repeating array.

Conveniently, each supply plena may have an elongate shape on a cross-section through the plenum parallel to the local plane of the external wall. For example, the elongate shape may be a racetrack shape. The elongate shape may have an aspect ratio in the range from 2:1 to 4:1. The supply plena may be arranged such that the length directions of their respective elongate shapes are parallel. The supply plena may be arranged such that the breadth directions of their respective elongate shapes are parallel to the local direction of maximum gradient of the pressure of the working gas over the external wall. In this way, each supply plenum can be made to extend over an area of the external wall which experiences little variation in working gas pressure. This helps the convective heat transfer rate and film cooling blowing rate to be optimised at the plenum. The spacing between any one of the supply plena and its nearest neighbour supply plena may be equal to at most twice the breadth of its elongate shape. In this way, all points on the external wall can benefit from local convective heat transfer and film cooling. The spacing between any one of the supply plena and its nearest neighbour supply plena may be equal to at least the breadth of its elongate shape. Dividing walls between the plena can thus have a minimum thickness which is castable and provides sufficient strength.

Rather than being racetrack shaped, the supply plena may have other shapes on cross-sections through the plena parallel to the local plane of the external wall. For example, each plenum may have a triangular, square, rectangular, circular, elliptical or hexagonal shape.

The component may be a turbine blade or a vane. In particular, the supply plena can be located at the aerofoil surface of the blade or vane, for example at or adjacent the leading edge of the aerofoil surface. However, additionally or alternatively, the supply plena can be located at platform surfaces of the blade or vane.

The component may be a shroud or shroud segment.

Embodiments of the invention will now be described by way of example with reference to the accompanying drawings in which: Figure 1 shows an isometric view of a typical single stage cooled turbine; Figure 2 shows a schematic longitudinal cross-section through a ducted fan gas turbine engine; Figure 3 shows a schematic transverse cross-section through a high-pressure turbine rotor blade; Figure 4 shows a schematic transverse cross-section through a high-pressure turbine nozzle guide vane; and Figures 5 to 16 are schematic diagrams of various possibilities for regular arrays of supply plena which supply metered cooling air to film cooling holes formed in the external wall of an aerofoil.

With reference to Figure 1, a ducted fan gas turbine engine incorporating the invention is generally indicated at 10 and 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, an 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 dud 22 and a bypass exhaust nozzle 23.

During operation, 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 13 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.

Figure 3 shows a schematic transverse cross-section through a rotor blade 40 of the high-pressure turbine 16, and Figure 4 shows a schematic transverse cross-section through an NGV 41 of the high-pressure turbine. Both the aerofoils have a central cavity 42 which accepts cooling air from one or more inlets (not shown). Large 43 and small 44 supply plena line portions of the inner surfaces of the external wall of the respective aerofoil. The plena are fed cooling air from the respective central cavity 42 via impingement holes 45. These holes produce impingement jets on the inner surfaces of the external walls to enhance convective cooling. They also meter the cooling air so that a controlled pressure drop is formed between the central cavity and the respective plenum. The supply plena then supply the metered cooling air to film cooling holes 46 formed in the external wall of the respective aerofoil, the air blowing through the cooling holes to form a cooling film on the outer surface of the external wall.

In the regions R encircled by dashed lines in Figures 3 and 4, the supply plena 44 are particularly small, each having only one or two impingement holes 45 and only one or two film cooling holes 46. Typical locations for these small plena 44 are the early suctions surface and early pressure surface of the aerofoil.

The present invention aims at improving cooling efficiencies in gas turbine engine components by enabling improvements in both internal convective cooling and film cooling effectiveness and coverage. For example, in the context of an aerofoil component, this aim can be achieved by setting cooling air pressures at the different levels locally around the aerofoil. The desired pressure level is where the film cooling blowing rate and impingement jet pressure ratio are collectively optimised, i.e. they deliver the required cooling effectiveness for the minimum amount of cooling flow or more correctly the lowest stage efficiency loss.

Large supply plena supplying cooling air to film cooling holes may not allow the pressure to be kept at the desired level because the plenum length covers a large surface area over which there is a range of local boundary conditions, and one value will not satisfy all.

However, the present invention allows selected parts of the surface to be broken up into a larger number of smaller regions over which the boundary conditions are changing less. As a result, impingement and film cooling geometries can be configured to more closely suit the local boundary conditions. In particular, the small supply plena 44 have only one or two metering feeds (impingement holes) 45 and one or two film cooling holes 46 associated with them. Thus impingement hole (jet) diameter(s) and film cooling hole diameter(s) can be more easily varied to suit local conditions.

The orientations of the small supply plena 44 can also be changed more easily to suit changes in external boundary conditions. For example, the plena can be shaped to have a shod dimension in the direction of large pressure gradients and a longer dimension in the direction of low pressure gradients.

The small supply plena 44 can also help to reduce the thermal stresses experienced by the ceramic core pedestals that during casting form the metering feeds linking the central cavity 42 to the individual plena at the aerofoil external wall. In regions where the aerofoil wall curvature is high, such as in the vicinity of the aerofoil leading edge, and where core breakage problems are particularly acute, the small supply plena effectively reduce local curvatures to alleviate stress problems and core breakages. Also the many separating walls between the small supply plena allow the heat from the hotter external wall of the aerofoil to be conducted through to the cooler inner wall between the plena and the central cavity, thus reducing the thermal fight between the two walls, and improving the thermal fatigue life of the aerofoil.

Each small supply plena 44 may have an elongate shape, such as a racetrack shape, on a cross-section through the plena parallel to the local plane of the external wall. Typically, the elongate shape has an aspect ratio in the range from 2:1 to 4:1. In each plenum, the location of impingement of the or each impingement jet on the inner surface of the external wall can be spaced from the location(s) on that wall of the entrance(s) to the film cooling hole(s).

The small supply plena 44 may be arranged in a regular repeating array extending over the external wall. Figures 5 to 16 are schematic diagrams of various possibilities for such regular arrays, each diagram showing the direction of hot working gas flow, and also showing for each plenum of the respective array: the racetrack cross-sectional shape of the plena 44, the location 47 of impingement of the or each impingement jet, the film cooling hole(s) 46, and the cooling air film trajectories 48 from the film cooling hole(s). In the arrays, the spacing between each supply plena and its nearest neighbour plena is generally no more than twice the breadth of the racetrack cross-sectional shape. This helps to ensure adequate cooling film coverage and internal heat transfer.

The array of Figure 5 has the plena arranged in a horizontal staggered arrangement with the films aligned with the horizontal. The film coverage protects the gap between plenum chambers and the high levels of internal heat transfer under the impingement jets and in the film cooling holes cools the regions not covered by the film trajectories.

The array of Figure 6 has the plena arranged in a vertical staggered orientation with the film trajectories aligned with the horizontal. The film coverage partially covers the gaps between the plena and washes over each neighbouring downstream plenum and the entrance to its film cooling hole. There are some areas that have little or no film cooling protection, but the use of shaped film cooling holes can improve this situation.

The array of Figure 7 has the plena arranged in an in-line horizontal orientation with the film cooling holes angled at approximately 40° to the horizontal. The film coverage protects the gap between the plena, while the high heat transfer rates within the plena and the film cooling holes convectively cools the zones between the film trajectories.

The array of Figure 8 has the plena arranged in a vertical staggered orientation with the film cooling holes angled at approximately 45° to the horizontal. As seen previously, the film coverage protects the gaps between plena with the film trajectories oriented such that they fill the zones between the impingement jets and the film cooling holes.

The array of Figure 9 has the plena arranged in a vertical partially staggered arrangement with two film cooling holes per plenum aligned with the horizontal. The film coverage is good over both the plena and the gaps between them. The film trajectories avoid neighbouring film cooling hole exits.

The array of Figure 10 has the plena arranged in a vertical partially staggered arrangement with two film cooling holes per plenum angled at approximately 45° to the horizontal. Again the film coverage is good over both the plena and the gaps between them, although some film trajectories flow over neighbouring film cooling hole exits.

The array of Figure 11 has the plena arranged in an in-line horizontal orientation with two film cooling holes per plenum angled at approximately +45° and -45° to the horizontal. The film coverage protects the gap between the plena, and high levels of internal heat transfer under the impingement jets and in the film cooling holes cools the zones between the film trajectories.

The array of Figure 12 has the plena arranged in both horizontal and vertical in-line configurations with one and two film cooling holes per plenum angled at approximately +45° and -45° to the horizontal. The film coverage protects the gap between the plena and high levels of internal heat transfer under the impingement jets and in the film cooling holes cools the zones between the film trajectories.

The arrays of Figures 13 to 16 show further arrangements, but now with each plenum having two impingement jets. In Figure 13 the plena are arranged in a vertical staggered orientation with the film cooling holes (one per plenum) aligned to the horizontal. In In Figure 14 the plena are arranged in a vertical staggered orientation with the film cooling holes (one per plenum) angled at approximately 45° to the horizontal. In Figure 15 the plena are arranged in a partially vertical staggered orientation with the film cooling holes (two per plenum) angled at approximately +45° and -45° to the horizontal. In Figure 16 the plena are arranged in a partially horizontally staggered orientation with the film cooling holes (two per plenum) angled at approximately +45° and -45° to the horizontal.

The arrays of Figures 5 to 16 are by no means exhaustive. For example, the long directions of the plena racetrack cross-sectional shapes could be at an angle of e.g. 45° to the horizontal instead of being vertically or horizontally aligned.

"Soluble core technology" is a known approach of producing ceramic cores that can be used to form cast components having supply plena arrangements of the type described above. In the approach, inserts are placed in the ceramic core die, the inserts being removed by dissolution once the ceramic core has been injected. However, there are other ways of producing suitable final core shapes, such as assembling a single core from a number of individual simpler cores. Another approach is to produce cores by 3D printing where, without the need for tooling, layers of ceramic are laminated together and subsequently heated at pressure to create a single homogeneous core.

Core ties may be used to link plena together in order to add strength to the fragile cores.

These ties can then be removed before the cores are loaded into the wax die so that there are no links between plena in the cast part. If they are left in place, then passages linking the plena are formed in the cast part. As these passages will leak cooling air between the plena, the passages should be sized (i.e. made sufficiently small) such that the cooling air in linked plena can be at different pressures.

A further manufacturing method that can be used to form the component directly is Direct Laser Deposition (DLD). This technology produces the component by solidifying or sintering thin layers of metallic powder using a focused laser. Each layer attaches to the previous layer of material allowing the component to be built layer-by-layer with all its heat transfer features, including impingement holes and film cooling holes. The component can thus be manufactured without the need for ceramic cores, core dies, wax dies or hole drilling equipment.

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. For example, as well as hot section NGV and blade aerofoils, the supply plena arrangements could beneficially be incorporated into NGV platforms, blade platforms and shroud segments. As another example, the plena may have shapes other than a racetrack shape on cross-sections through the plena parallel to the local plane of the external wall. 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 (13)

1. CLAIMS1. A component of a gas turbine engine, the component including: an external wall (40) which, in use, is exposed on the outer 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 the supply plena, and the supply plena supplying the metered cooling air to the cooling holes; and wherein each supply plenum is fed by just one or two of the metering feeds, and supplies the metered cooling air to just one or two of the cooling holes.
2. 2. A component according to claim 1, wherein the supply plena are partially defined by the inner surface (46) of the external wall.
3. 3. 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 the inner surface.
4. 4. A component according to any one of the previous claims which further includes an internal cavity (44) in flow series between the air inlet arrangement and the supply plena, the metering feeds feeding the cooling air from the internal cavity to the supply plena.
5. 5. A component according to any one of the previous claims, wherein the supply plena are arranged in a regular repeating array.
6. 6. A component according to any one of the previous claims, wherein each supply plena has an elongate shape on a cross-section through the plenum parallel to the local plane of the external wall.
7. 7. A component according to claim 6, wherein the supply plena are arranged such that the length directions of their respective elongate shapes are parallel.
8. 8. A component according to claim 6 or 7, wherein the supply plena are arranged such that the breadth directions of their respective elongate shapes are parallel to the local direction of maximum gradient of the pressure of the working gas over the external wall.
9. 9. A component according to any one of claims 6 to 8, wherein the elongate shape has an aspect ratio in the range from 2:1 to 4:1.
10. 10. A component according to any one of claims 6 to 8, wherein the spacing between any one of the supply plena and its nearest neighbour supply plena is equal to at most twice the breadth of its elongate shape.
11. 11. A component according to any one of the previous claims which is a turbine blade or a vane.
12. 12. A gas turbine engine having one or more components according to any one of the previous claims.
13. 13. A component of a gas turbine engine as any one herein described with reference to and as shown in Figures 3 to 16.