GB2492374A

GB2492374A - Gas turbine engine impingement cooling

Info

Publication number: GB2492374A
Application number: GB201111105A
Authority: GB
Inventors: Mark Timothy Mitchell
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2011-06-30
Filing date: 2011-06-30
Publication date: 2013-01-02
Also published as: GB201111105D0

Abstract

A gas turbine engine component has first and second walls 52 that define a passageway 61 for the flow of a fluid C. The second wall defines at least one impingement hole 70 for the flow of an impingement fluid A. The second wall comprises a deflector wall 80 around at least a part of the impingement hole and extending from the second wall towards the first wall. The deflector wall may extend between 25% and 95%, and preferably 75%, of the distance between the walls. The deflector wall may surround the entire impingement hole, or may surround a portion, preferably at least 90Â°, of the hole. An inner surface of the deflector wall may be continuous with the impingement hole, or spaced there from. A plurality of impingement holes may be provided, and downstream (relative to fluid flow C) holes may extend further towards the first wall. The component may be an aerofoil (blade) or combustor wall.

Description

IMPINGEMENT COOLING ARRANGEMENT

The present disclosure relates to a cooling arrangement for use on components within gas turbine engines. The impingement cooling arrangement may be particularly, although not exclusively, useful when implemented on aerofoil and combustor tile components in a gas turbine engine.

It s usual that components ri hot stages of a gas turbine engine will require cooling during normal operating conditions. Cooling is particularly required for components functioning in the hottest turbine stage of a gas turbine engine.

Conventionally, in the high-pressure turbine stage the gas stream or working fluid passing through the gas turbine engine can reach temperatures in excess of 1600°C. This temperature can be greater than the melting point of the materials, typically nickel-based alloys, used to manufacture components of the turbine.

Turbine rotor blades and nozzle guide vanes are components which, without cooling, would have short in-service lives. Furthermore, adequate cooling and temperature control helps to maximise component life.

Conventionally cooling is achieved by utilising a portion of the air drawn into the engine during operation. Usually this cooling air is bled from the high-pressure compressor in the core of the gas turbine engine. Although relatively cool in comparison to the temperature of the working fluid at the hottest turbine stage of the engine, the cooling air may be at a temperature in excess of 700°C.

This is due to the work imparted by the compressor stages of the engine. The cooling air is utilised, with various techniques, to cool a range of components within the engine.

One technique to cool components employs striking cooling air against a surface of the component requiring cooling. Known in the art as impingement cooling, this technique requires directing a discrete jet of cooling air onto the surface of the hot component. This arrangement has particular application for nozzle guide vanes and rotor blades within the turbine stages of a gas turbine engine. The present cooling arrangement is equally applicable to combustion components of a gas turbine engine.

This striking action of the cooling air, against the surface, transfers heat from the component into the cooling air. The utilised or spent cooling air is then discarded through openings in the outer surface of the component; usually, this is into the working fluid passing over the component. This is because the component will conventionally be operating in the hot gas stream. Additionally, the spent cooling air may contribute to other cooling techniques, such as film cooling, which work in concert with impingement cooling to make up the total cooling required by the component.

The cooling air is usually directed at the surface in the form of a jet, or a pluraNty of jets. The jets are farmed by openings defined by a wall offset from the surface to be cooled. The cooling air is fed under pressure to one side of the wall and then flows through the openings in jet form to strike the surface. The spent cooling air then flows, in a transverse direction to the jets of air, between the wall and the surface towards the openings in the outer surface of the component. A problem can arise wherein the transverse flow can be responsible for lessening the effective strength, and thus the cooling effect, of the impingement jets against the surface.

Bleeding air, especially from the core of the gas turbine engine for cooling purposes reduces the efficiency of the engine despite, as described above, the spent cooling air being returned to the working fluid. The reasons for this are twofold: firstly, a cycle penalty is incurred by not using the bled cooling air for combustion and so reducing the amount of energy, and thus work, transferable to the turbine; secondly, aerodynamic losses are incurred when the spent cooling air is returned to the gas stream through the openings in the component.

The necessity for cooling and cooling air presents a design challenge for gas turbine engines when striving for increased efficiency and therefore a reduction in specific fuel consumption.

If a more effective rate of cooling can be achieved per unit of cooling air, then it follows that it may be preferable for a commensurate reduction the amount of cooling air bled from the core. A more effective rate of cooling may then improve the gas turbine engine cycle efficiency.

Alternatively, a higher available cooling capacity may allow for an increase in the temperature tolerance of components, such as turbine nozzle guide vanes.

The efficiency of the work cycle of the gas turbine engine can be increased as a result of an increase in the working fluid temperature at the turbine inlet.

Additionally, an increase in turbine inlet temperatures may lead to reductions of pollutants in the gas stream and therefore in the engine exhaust. However, these efficiency design factors must be balanced against a reduction in component life due to the higher operating temperatures.

In accordance with the present disclosure a component for a gas turbine engine is provided; the component comprising a first wall and a second wall defining a passageway for a flow of a fluid (C), the second wall defining at least one impingement hole for the flow of an impingement fluid (A), the second wall comprises a deflector wall around at least a part of the impingement hole and extending from the second wall towards the first wall.

The deflector wall may extend from the second wall towards the first wall to a distance of between (twenty-five percent) and (ninety-five percent) of the total distance between the first wall and the second wall.

The deflector wall may extend from the second wall towards the first wall to a distance of (seventy-five percent) of the total distance between the first wall and the second wall.

The deflector wall may surround the entire impingement hole.

Alternatively, the impingement hole may have a central axis therethrough and the deflector wall may extend about a portion of the impingement hole defined by an angle (a) about the central axis.

The angle (a) may be bisected by a direction of the flow of the fluid (C).

The angle (a) may be at least 90°.

The deflector wall may form a continuous inner surface with the impingement hole.

Alternatively, the deflector wall may be located a distance away from the impingement hole.

Alternatively, the deflector wall may occupy a position between adjoined to the impingement hole, and a distance of up to three times the diameter of the impingement hole away from the impingement hole.

The impingement hole may be one of an array of impingement holes.

The array of impingement holes may include a second impingement hole, defined by the second wall, located in an upstream direction of the flow of the fluid (C) from the impingement hole; the second wall comprising a second deflector wall around at least a part of the second impingement hole and extending from the second wall towards the first wall; and the deflector wall of the impingement hole extending towards the first wall by a greater distance than the second deflector wall extends towards the first wall.

The array of impingement holes may have each impingement hole including the deflection wall.

The component may comprise the array of impingement holes having deflector walls, the deflector walls extending progressively further towards the second wall in a downstream direction of the flow of the fluid (C).

The component may be any one of a group comprising an aerofoil or a combustor wall.

Accordingly, embodiments of the present disclosure will now be described by way of example with reference to the following drawings in which: Figure 1 is a longitudinal section of a gas turbine engine of the type known in the art as a turbofan or ducted bypass fan engine and; Figure 2 is a generalised section of a turbine stage from a gas turbine engine in which the disclosed impingement cooing arrangement is implemented; Figure 3 is a related art impingement cooling system shown in illustrative form; Figure 4 is a view through Section X-X of an aerofoil in which the disclosed impingement cooing arrangement is implemented; Figure 5a is an isometric part-section through an aerofoil comprising a preferred embodiment of the present impingement cooling arrangement; Figure 5b is an illustrated schematic of the preferred embodiment as described with respect to Figure 5a; Figure 5c is a view of Section Y-Y through the aerofoil as indicated in Figure 5a;and Figures 6a and 6b are views of Section Y-Y showing alternative embodiments of the present impingement cooling arrangement.

With reference to Figure 1, a gas turbine engine power plant 10, of the type known in the art as a turbofan or ducted bypass fan engine, is shown. The gas turbine engine 10, having a principle rotational axis 11, is shown in section comprising a low pressure fan 13, intermediate pressure compressor 14, a high pressure compressor 15, a combustion chamber 16, a high pressure turbine 17, an intermediate pressure turbine 18, a low pressure turbine 19 and a core engine exhaust duct 20. A nacelle 21 generally surrounds the engine and with a nose cone 38 defines a ducted inlet 12. The nacelle 21 also defines a bypass duct 22 and a bypass exhaust 23.

Conventionally gas turbine engines operate by drawing air through the ducted inlet 12 by way of the rotating low pressure compressor or fan 13. A significant portion of the air drawn through the fan 13 is discharged directly through the bypass duct 22 which connects with the bypass exhaust 23 contained within the nacelle 21. The remaining air entering the gas turbine engine passes through, what is commonly termed, the core of the engine entering via its intermediate pressure compressor 14 where the air is compressed. A further stage of compression is provided by the high pressure compressor 15.

Air exiting the high pressure compressor enters the combustion chamber 16 where fuel is added and the resulting mixture of air and fuel is combusted creating a high temperature working fluid. The working fluid exits the combustion is chamber 16 and enters the high pressure turbine 17. The high pressure turbine 17 is drivingly connected to the high pressure compressor 15 is via a shaft (not referenced). The expanding working fluid exits the high pressure turbine into two further turbine stages: the intermediate pressure turbine 18 and the low pressure turbine 19, which are drivingly connected to the intermediate pressure compressor 15 and the low pressure compressor 16 respectively via shafts (not referenced). The working fluid or combusted mixture exits the low pressure turbine 19 into the core engine exhaust duct 20 ultimately exiting the gas turbine engine.

Referring now to Figure 2, where a turbine stage 40 is shown in generalised section and may be from any one of the turbines 17, 18, 19 of the gas turbine engine. Conventionally, the turbine stage 40 includes an annular array of stationary nozzle guide vanes 41, an annular array of turbine rotor blades 42 fixed to a turbine rotor disc 43 able to rotate about the principle axis 11, an inlet passageway 45 and an exit passageway 46. The nozzle guide vanes 41 comprises an aerofoil 50 between its radially inner and outer extremities.

The working fluid passes through the inlet passage way 45 and is directed on to the turbine rotor blades 42 at the appropriate angle by the nozzle guide vanes 41. The working fluid acts on the turbine rotor blades 42 which react and drive the turbine rotor disc 43 assembly. The working fluid exits the turbine stage through exit passageway 46.

Extracting work efficiently from the high-temperature working fluid in the high pressure turbine 17 requires the array of turbine nozzle guide vanes 41 and the corresponding turbine rotor blades 42 to be immediately downstream of the combustion chamber 16. Conversely the efficiency of the gas turbine engine is reduced as a large amount of cooling is required to maintain manageable operating temperatures of the nozzle guide vanes 41 and the turbine rotor blades 42. The high pressure turbine 17 stage will require the most cooling while the intermediate pressure 18 and low pressure turbine 19 stages require progressively less cooling as the working fluid flowing through the engine reduces in temperature.

Typically the cooling is provided by air bled from the high pressure compressor 15 by-passing the combustion chamber 16 and therefore remaining relatively cool. Primarily the cooling air is used for internal forced convection, impingement and film cooling within components of the turbine stages. Bleeding cooling air from the high pressure compressor 15 adversely affects the efficiency of the engine as the portion of intake air contributing to the working fluid is reduced. Using cooling air more effectively creates an opportunity to reduce the quantity of bleed air or increase the turbine inlet temperature which, as discussed above, can improve gas turbine engine operating cycle efficiency.

An established, but not exclusive, means of impingement cooling a gas turbine engine component 100 is shown in illustrative form in Figure 3. As known in the art, impingement cooling is achieved by striking a surface 101 of the component 100 to be cooled with an impingement jet (A) of air; the impingement of the air on the surface 101 being an effective means to transfer heat from the component 100 to the air. In a known arrangement, a wall 103 neighbours the surface 101 defining a passageway 104 therewith and an impingement hole 106.

A cavity 105 is isolated from the passageway 104 by the wall 103.

High pressure air is fed in to the cavity 105 and at least a potion of the air is forced through the impingement hole 106 forming the impingement jet (A). The impingement hole 106 also serves to direct the impingement jet (A) on to the surface 101. Utilised coolant in the form of flow B is discarded through the passageway 104. Generally, efficient heat transfer will be achieved if the impingement jet (A) of air is directed perpendicularly at the surface 101, or, at least at a relatively high angle of incidence.

Often the impingement hole 106 is one of an array of similar holes in order to realise the large amount of cooling necessary for elements of gas turbine engines. A number of array arrangements exist to meet differing cooling objectives in components.

In an impingement coaling arrangement a problem can be caused by the spent coolant flow B and/or any other flow passing through the passageway 104.

Across an array of impingement holes 106 the cumulative or additive effect of the flows (B) from each hole results in a flow C travelling in a transverse direction towards an ultimate exit point, which is not shown, from the passageway 104.

Flow C is often strong enough to act in shear on the impingement jet (A) and reduce impingement on to the surface 101, or in an extreme case, destroy the jet entirely. This problem is particularly apparent for those impingement jets which are closer to the ultimate exit point or in the downstream direction of flow C. One solution entails increasing coolant to the impingement holes 106 to create stronger impingement jets (A).

Referring now to Figure 4 showing a Section X-X through the aerofoil 50 from the array of nozzle guide vanes 41 indicated in Figure 2. It should be understood by the skilled artisan that this arrangement could equally apply an element of the array of turbine blades 42 or other double-walled structure such as a combustor. The architecture of the aerofoil 50 is known in the art and comprises a suction wall 51 and an inner wall 52 defining an inner passageway 61 analogous with the passageway 104 present in Figure 3. A pressure wall 53 meets with the suction wall 51 at a leading edge 56 and a trailing edge 55 of the aerofoil 50. A web 53 further divides the internal volume of the aerofoil 50 into a rear cavity 62 and a front cavity 63, extending from the pressure wall 53 to the suction wall 51. A plurality of impingement holes 70, comparable with the impingement hole 106 described in Figure 3, are defined by the inner wall 52.

The pressure wall 53 defines an array of film cooling holes 72 and in conjunction with the suction wall 51 defines an exit slat 71 in the trailing edge 55. The exit slot 71 may be considered analogous with the ultimate exit point of the passageway 104 in Figure 3.

The arrangement of the aerofoil 50 is the implementation of the art as illustrated in Figure 3. During engine operation the working fluid will impart heat into the aerofoil 50 and cooling is therefore required to maintain operating temperature and structural integrity. Relatively cool air at about 700°C is injected in a radial direction from one or both ends of the front cavity 63 and the rear cavity 62. Typically the coolant is injected at around 3,800kPa, a marginally higher pressure than the working fluid passing aver the pressure wall 53 and suction wall 51; nevertheless, this small margin is critical to ensure coolant flow through the aerofoil 50 to maintain component integrity.

Coolant entering the front cavity 63 is discharged through the film cooling holes 72 forming a cooling film on the surface of the suction wall 51 and pressure waIl 53. Similarly a portion of cooling air flow entering the rear cavity 62 contributes to the film cooling on the surface of the pressure wall 53. The remaining portion of the coolant entering the rear cavity 62 is utilised for impingement cooling of the suction wall 51 in the manner illustrated in Figure 3.

In this case, coolant passing through the impingement holes 70 and striking on the internal surface of the suction wall 51 is analogous to the impingement jet (A) of air in Figure 3. The impingement of the coolant on this internal surface conducts and removes thermal energy from the suction wall 51 and so contributes to maintaining structural integrity. Utilised coolant is discharged in a rearward manner through the inner passageway 61 and generally parallel to the suction wall 51. Discharge of the coolant through the exit slot 71 contributes to the cooling of the trailing edge 55 of the aerofoil 50.

The flow of coolant through the inner passageway 61 towards the trailing edge 55 is analogous to flow C in Figure 3. The impingement cooling utilised in the aerofoil 50 is effective in removing heat from the suction wall 51; however, the rearward flow of utilised coolant, in a similar manner to flow C in Figure 3, can reduce the effectiveness of the impingement holes 70 which are closer to the trailing edge 55.

Figure 5a is a preferred embodiment of the present disclosure showing an isometric part-section through the aerofoil 50 having an array of impingement holes 70 one of which is an impingement hole 85. A deflector wall 80 surrounds the impingement hole 85 and forms a continuous inner surface 81 with the impingement hole 85. The continuous inner surface 81 may be in the form of a cylindrical bore, typically shaped by the hole forming process. While it is typical for the impingement hole 85 to be cylindrical in shape, it should be appreciated that the impingement hole 85 could be an opening of any shape; for instance, an elliptical shaped opening. This may be driven by the manufacturing method utilised to form the impingement hole 85. The deflector wall 80 projects from the cooler inner wall 52 towards the hotter suction wall 51 of the aerofoil 50 on to which the impingement jet (A) strikes. The length of the projection of the deflector wall 80 may be expressed as a percentage of the overall distance between the inner wall 52 and the suction wall 51; or, alternatively, as a ratio between the diameter of the impingement hole 85 and length of the projection of the deflector wall 80.

The applicant believes a distance of seventy-five percent is particularly useful although percentage distances greater than twenty-five percent are useful. The maximum percentage distance is approximately ninety percent although this can vary dependent on mass flow of the impingement jet (A), where there is less flow then a greater distance of up to ninety-five percent of the overall distance may be preferable.

Figures 5b and Sc are illustrated schematics of the preferred embodiment of the aerofoil 50 as described with respect to Figure 5a. Figure Sb is a view in the direction of Z in Figure 5a. Figure Sc is a view of Section Y-Y in Figure Sa.

Figure Sb and Figure Sc depict the coolant flow in the embodiment of the gas turbine engine during operation.

As described above, the impingement jet (A) is formed by the impingement hole 85 fed by air supplied to the rear cavity 62. In the preferred embodiment the deflector wall 80 protects the impingement jet (A) from the transverse flow C as it travels towards the exit slot 71. Because of the deflector wall 80, the transverse flow C has less distance to disrupt, or destroy, the flow profile of the impingement jet (A).

It may be additionally advantageous that favourable flow conditions created by the deflection wall 80 divert and encourage flow C into areas of the inner passageway 61 offering less resistance. This flow condition is shown in Figure Sc. As described above, the transverse flow C is cumulatively augmented by spent impingement coolant flow B as flow C travels towards the exit slot 71. In some aerofoil arrangements the flow C might be from a source other than that derived from impingement cooling flows.

Advantageously, heat transfer may be potentially improved as the velocity of the jet of air A is locally increased at the point of impingement on the surface of suction wall 52 where it assumes the form of flow B. Referring again to Figure 5a, where it should be appreciated that an array of impingement holes 70 may comprise a combination of holes both with, and without, a deflector wall 80. For example, a deflector wall 80 may not be required for an impingement hole 85 where the transverse flow C in the inner passageway 61 is weak or even non-existent. This situation is likely to occur near impingement holes 70 which are located furthest or upstream from the exit slot 71. The preferred embodiment of the disclosure may, however, include a deflector wall 80 for each impingement hole 85. A further alternative arrangement combating the variability of the strength of the flow C, as described above, is the array of impingement holes 70 having deflector walls 80 of differing projection lengths. Thus in the case of a weaker flow C further upstream from the exit slot 71, the impingement hole 85 may have a smaller projection length of deflector wall 80 than an impingement hole 85 closer to the exit slot 71.

Further embodiments of the present disclosure are depicted in Figures 6a- 6b. Figure 6a shows that the deflection wall 80 does not entirely surround the impingement hole 85; here, the deflection wall 80, occupies only a portion of the periphery of the opening. The portion of the opening protected by the deflection wall 80 may be defined in terms of an angle a and oriented relative to the direction of the flow C. The angle a may be measured about a central axis 90 defined by the impingement hole 85 and equally bisected by flow C travelling downstream or towards the impingement hole 85. Angle a may range from approximately 90° to 360° as defined in the embodiment depicted in Figure 6a.

The applicant believes that an alternative embodiment of the cooling arrangement is where the deflector wall 80 is arranged such that it projects from the inner wall 52 and does not form the a continuous inner surface 81 with the impingement hole 85 as depicted in Figure 5a; rather, the deflector wall 80 is set back from the periphery of the impingement hole 85. This is shown in Figure 6b, where again only a portion of the opening is protected by the deflector wall 80, defined according to the angle a and oriented relative to the direction of the flow C. Figure 6b also shows a range may apply for angle a, as for the embodiment depicted in Figure 6a, including an embodiment where the deflector waIl 80 is set back and surrounds the entirety of the impingement hole 85. The applicant believes that the deflector wall 80 in this embodiment is preferably within a distance from the periphery of the impingement hole 85, the distance being three times the impingement hole 85 diameter.

Claims

<claim-text>CLAIMS1. A gas turbine engine component; the component comprising a first wall and a second wall defining a passageway for a flow of a fluid (C), the second wall defining at least one impingement hole for the flow of an impingement fluid (A), the second wall comprises a deflector wall around at east a part of the impingement hole and extending from the second wall towards the first wall.</claim-text> <claim-text>2. A gas turbine engine component as claimed in claim 1 wherein the deflector wall extends from the second wall towards the first wall to a distance of between twenty-five percent and ninety-five percent of the distance between the first wall and the second wall.</claim-text> <claim-text>3. A gas turbine engine component as claimed in claim 1 wherein the deflector wall extends from the second wall towards the first wall to a distance of approximately seventy-five percent of the distance between the first wall and the second wall.</claim-text> <claim-text>4. A gas turbine engine component as claimed in any one of claims I -3 wherein the deflector wall surrounds the entire impingement hole.</claim-text> <claim-text>5. A gas turbine engine component as claimed in any one of claims 1 -3 wherein the impingement hole has a central axis therethrough and wherein the deflector wall extends about a portion of the impingement hole defined by an angle (a) about the central axis.</claim-text> <claim-text>6. A gas turbine engine component as claimed in claim 5 wherein the angle (a) is bisected by a direction of the flow of the fluid (C).</claim-text> <claim-text>7. A gas turbine engine component as claimed in claim 5 wherein the angle (a) is at least 900.</claim-text> <claim-text>8. A gas turbine engine component as claimed in any one of claims 1 -7 wherein the deflector wall forms a continuous inner surface with the impingement hole.</claim-text> <claim-text>9. A gas turbine engine component as claimed in any one of claims 1 -7 wherein the deflector wall is located a distance away from the impingement hole.</claim-text> <claim-text>10. A gas turbine engine component as claimed in any one of claims 1 -7 wherein the deflector wall occupies a position between the impingement hole and a distance of up to three times the diameter of the impingement hole away from the impingement hole.</claim-text> <claim-text>11. A gas turbine engine component as claimed in any one of claims 1 -10 where the impingement hole is one or more of an array of impingement holes.</claim-text> <claim-text>12. A gas turbine engine component as claimed in claim 11 wherein the array of impingement holes includes a second impingement hole, defined by the second wall, located in an upstream direction of the flow of the fluid (C) from the impingement hole; the second wall comprising a second deflector wall around at least a part of the second impingement hole and extending from the second wall towards the first wall; and the deflector wall of the (first) impingement hole extending towards the first wall by a greater distance than the second deflector wall extends towards the first wall.is 13. A gas turbine engine component as claimed in claim 11 wherein a deflector wall surrounds at least a part of each impingement hole.14. A gas turbine engine component as claimed in claim 11 wherein the component comprises the array of impingement holes having deflector walls, the deflector walls extending progressively further towards the second wall in a downstream direction of the flow of the fluid (C).15. A gas turbine engine component as claimed in any one of claims 1 -14 wherein the component is any one of a group comprising an aerofoil or a combustor wall.16. A gas turbine engine component as described with reference to Figures 4 -6.</claim-text>