WO2010123415A1

WO2010123415A1 - An annular gas turbine housing component and a gas turbine comprising the component

Info

Publication number: WO2010123415A1
Application number: PCT/SE2009/000219
Authority: WO
Inventors: Dennis Jacobsson
Original assignee: Volvo Aero Corporation
Priority date: 2009-04-24
Filing date: 2009-04-24
Publication date: 2010-10-28
Also published as: US20120141256A1; EP2422054A1

Abstract

The invention relates to an annular gas turbine housing component (15) comprising a plurality of circumferentially spaced first channels (16) for conveying a first gas flow in a first direction and a plurality of circumferentially spaced second channels (17) for conveying a second gas flow in a second direction across the first direction in a crossing position (18) of the first and second channels.

Description

An annular gas -turbine housing component: and a gas turbine comprising the component

FIELD OF THE INVENTION

The present invention relates to an annular gas turbine housing component according to the preamble of claim 1. The invention is further directed to a gas turbine engine, and especially to an aircraft engine,^' comprising the component. Thus, the invention is especially directed to a jet engine.

Jet engine is meant to include various types of engines, which admit air at relatively low velocity, heat it by combustion and shoot it out at a much higher velocity. Accommodated within the term jet engine are, for example, turbojet engines and turbofan engines. The invention will below be described for a turbofan engine, but may of course also be used for other engine types.

An aircraft gas turbine engine of the turbofan type generally comprises a forward fan and booster compressor, a middle core engine, and an aft low pressure power turbine. The core engine comprises a high pressure compressor, a combustor and a high pressure turbine in a serial relationship. The high pressure compressor and high pressure turbine of the core engine are interconnected by a high pressure shaft. The high- pressure compressor is rotatably driven to compress air entering the core engine to a relatively high pressure. This high pressure air is then mixed with fuel in the combustor and ignited to form a high energy gas stream. The gas stream flows aft and passes through the high- pressure turbine, rotatably driving it and the high pressure shaft which, in turn, rotatably drives the high pressure compressor.

The gas stream leaving the high pressure turbine is expanded through a second or low pressure turbine. The low pressure turbine rotatably drives the fan and booster compressor via a low pressure shaft. The low pressure shaft extends through the high pressure rotor. Most of the thrust produced is generated by the fan.

There is a desire for improved performance and fuel efficiency in turbofan and other turbine engines. Specific thrust may be increased and specific fuel consumption may be decreased by increasing the cycle pressure ratio (CPR) of the engine. One known way of increasing the CPR beyond today's level is to cool the incoming air between the compressor stages. The annular gas turbine housing component may be arranged between the low pressure compressor and the high pressure compressor and configured to define the primary gas flow channel through the gas turbine engine.

US 6,314,880 discloses a gas turbine engine provided with an intercooler connected to the core air flow and constructed to transport the core air downstream of the low pressure compressor to the bypass air passage and back to the core air passage upstream of the high pressure compressor. More specifically, the air flow from the low pressure compressor is conveyed in a loop via circumferentially spaced channels provided in the bypass air passage, which channels convey the core air in the opposite direction in relation to the bypass air •flow, back to the core air passage. A diffuser is provided for directing the core air flow to the circumferentially spaced channels in the bypass air passage and a collector is provided for collecting the air flow from the circumferentially spaced channels in the bypass air passage before entering the high pressure compressor. The diffuser and the collector are arranged in a crossing relationship so that the core air to the intercooler crosses the core air from the intercooler.

SUMMARY OF THE INVENTION

An object of the invention is to achieve a gas turbine housing component comprising a wall structure, especially for application between compressor stages, which component is more rigid than prior art solutions and which creates conditions for low flow-losses. The improved rigidity creates conditions for an improved load-carrying ability. The component should further be cost-efficient in production while maintaining or improving its operational characteristics.

This object is achieved in a gas turbine housing component according to claim 1. Thus, it is achieved by an annular gas turbine housing component comprising a plurality of circumferentially spaced first channels for conveying a first gas flow in a first direction characterized in that the component comprises a plurality of circumferentially spaced second channels for conveying a second gas flow in a second direction across the first direction in a crossing position of the first and second channels.

Due to that the component comprises the plurality of circumferentially spaced second channels, which crosses the plurality of circumferentially spaced first channels, a rigid structure can be achieved.

Especially, a symmetric channel structure may be achieved with regard to a vertical plane along a center axis of the component and/or in a circumferential direction of the component. In order to achieve such a rigid structure, according to one example, a plurality of the first channels alternate with a plurality of the second channels in the circumferential direction.

According to a more specified example, every other channel in the circumferential direction is a first channel and every other channel in the. circumferential direction is a second channel at least along a portion of^■ the circumference. Preferably, every other channel in the circumferential direction is a first channel and every other channel in the circumferential direction is a second channel at least along a substantial part of the circumference. More preferably, every other channel in the circumferential direction is a first channel and every other channel in the circumferential direction is a second channel along the complete circumference.

Further, by arranging the plurality of circumferentially spaced second channels in a crossing relationship with the plurality of circumferentially spaced first channels, arrangements of sub-systems may be facilitated. For example, the radially spaced channels creates conditions for radially extending elements, such as power take off shafts, oil conduits etc.

According to a further example embodiment, the component is configured for a channel loop connecting said first channels to said second channels so that a gas flow from said first channels subsequently will flow through said second channels in gas turbine operation.

Such a design creates conditions for cooling the gas flow downstream of the first channels and upstream of the second channels. Preferably, a heat exchanger is positioned in said loop for cooling the gas flow.-

According to an example embodiment, at least a plurality of the first channels are configured so that there is an increase in their radial extension from a position upstream the crossing position in a gas flow direction to the crossing position. This diverging design of the first channels creates conditions for limiting a flow path area decrease, which would otherwise, be necessitated due to the crossing of the channels. In other words, a flow speed increase, which would otherwise be necessitated, is limited. Such an arrangement is for example desirable in case there is a heat exchanger provided in the loop according to the last-mentioned embodiment since the flow losses in the heat exchanger are correspondingly limited.

According to a further example embodiment, said plurality of first channels are configured so that a flow path area in said crossing position is at least as large as in a position upstream the crossing position in a gas flow direction. The "flow path area" is here defined as the total flow path area in all the first channels. Further, the "upstream position" is located at an exit of the upstream compressor stage in the gas turbine or a position inbetween the upstream compressor stage exit and the crossing position. A flow speed decrease can thereby be achieved. Such an arrangement is for example desirable in case there is a heat exchanger provided in the loop according to the last- β mentioned embodiment since the flow losses in the heat exchanger are correspondingly limited.

According to a further example embodiment, said plurality of first channels are configured so that a flow path area in a position upstream the crossing position in a gas flow direction is substantially the same as in said crossing position. Thus, the speed of the gas flow is substantially the same in the upstream position and the crossing position.

Other advantageous features and functions of various embodiments of the invention are set forth in the following description and in the dependent claims.

,BRIEF DESCRIPTION OF THE DRAWINGS

The invention will be explained below, with reference to the embodiment shown on the appended drawings, wherein FIG 1 is a schematic side view of an aircraft engine cut along a plane in parallel with the rotational axis of the engine,

FIG 2 and 3 are two schematic views of the flow channels in an intermediate housing component from figure

1, FIG 4 is a schematic, perspective view of a crossing point of the flow channels in figure 2 and 3, and FIG 5 and 6 shows two different embodiments of the flow channels in the intermediate housing component.

DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS OF THE

INVENTION

The invention will below be described for a turbofan gas turbine aircraft engine 1, which in figure 1 is circumscribed about an engine longitudinal central axis

2. The engine 1 comprises an outer casing or nacelle 3, an inner casing 4 (rotor) and an intermediate casing 5 which is concentric to the first two casings and 5 divides the gap between them into an inner primary gas channel 6 for the compression of air and a secondary channel 7 in which the engine bypass air flows. The casings are in turn made up of a plurality of components in the axial direction of the engine. Thus, 10 each of the gas channels 6,7 is annular in a cross section perpendicular to the engine longitudinal central axis 2.

The engine 1 comprises a fan 8 which receives ambient

15 air 9, a booster or low pressure compressor (LPC) 10 and a high pressure compressor (HPC) 11 arranged in the primary gas channel 6, a combustor 12 which mixes fuel with the air pressurized by the high pressure compressor

11 for generating combustion gases which flow downstream

20 through a high pressure turbine (HPT) 13 and a low pressure turbine (LPT) 14 from which the combustion gases are discharged from the engine.

A high pressure shaft joins the high pressure turbine 13

25. to the high pressure compressor 11 to substantially form a high pressure rotor. A low pressure shaft joins the low pressure turbine 14 to the low pressure compressor

10 to substantially form a low pressure rotor. The low pressure shaft is at least in part rotatably disposed

30 co-axially with and radially inwardly of the high pressure rotor.

An annular gas turbine housing component 15 is positioned between the low pressure compressor 10 and the high pressure compressor 11. The component 15 will now be further described with reference to figures 2-5. The arrows indicate the flow direction. The component 15 comprises a plurality of circumferentially spaced first channels 16 for conveying a first gas flow in a first direction and a plurality of circumferentially spaced second channels 17 for conveying a second gas flow in a second direction across the first direction in a crossing position 18 of the first and second channels. A gas flow from the low pressure compressor 10 will enter the first channels 16 and the gas flow from the second channels 17 will enter the high pressure compressor 11 in operation.

^■ More specifically, a plurality of the first channels 16 alternate with a plurality of the second channels 17 in the circumferential direction. Especially, every other channel in the circumferential direction is a first channel 16 and every other channel in the circumferential direction is a second channel 17 along the complete circumference.

The component 15 comprises a wall structure 19, which comprises a plurality of circumferentially spaced radial walls 20,120,220, which define said first channels 16 and second channels 17, see figure 3. The wall structure 19 is configured so that each of said radial walls 20,120,220 define one of said first channels 16 on one side and one of said second channels 17 on the other side in the circumferential direction at least along a portion of the circumference and preferably along the complete circumference. The wall structure is configured to be load-carrying. At- least a plurality of the first channels 16 are configured so that there is an increase 21 in their radial extension from a position 22 upstream the crossing position 18 in a gas flow direction to the crossing position 18. Each of said plurality of first channels 16 is configured so that a flow path area in said upstream position is substantially the same as in said crossing position. Alternatively, the flow path area in said upstream position may be smaller than in said crossing position. In any case, the gas flow path area is not drastically decreased in said crossing position in relation to said upstream position.

The component 15 comprises a circumferentially substantially continuous gas flow channel 23, which is divided into said plurality of circumferentially spaced first channels 16. An axial position of said channel division is about at the same axial position as the upstream position. In the shown embodiment, the upstream position 22 is defined by an axial position of said channel division.

The first channels 16 are configured for conveying the first gas flow substantially in an axial direction of the component 15 past the crossing point 18. More specifically, a center line 24 of -the continuous gas flow channel 23 in a radial direction in the upstream position 22 is positioned on substantially the same radial position as a center line 25 of said plurality of circumferentially spaced first channels 16 in said crossing position 18. Further, the plurality of second channels 17 are configured for conveying the second gas flow substantially in a radial direction of the component 15 past the crossing point 18. Accordingly, the first channels 16 are directed substantially perpendicular to the second channels 17 at the crossing point 18.

The component 15 creates conditions for a channel loop 26 connecting said first channels 16 to said second channels 17 so that a gas flow from said first channels

16 subsequently will flow through said second channels

17 in gas turbine operation. More specifically, a heat exchanger 27 is positioned in said loop for cooling the gas flow. The heat exchanger preferably comprises a cooling structure 28 with plurality of spaced tubes 29 for conveying the gas flow from the first channels 16 to the second channels 17. A plurality of cooling fins 30 are arranged between adjacent' cooling tubes 29 for improving the cooling performance .

In order to limit flow losses in the heat exchanger 27, the gas flow area should be substantially the same or may even be increased from an outlet of the upstream compressor (low pressure compressor) 10 to an inlet of the heat exchanger 27. Thus, the flow path area may be increased from the outlet of the upstream compressor to the crossing position 18 and preferably also downstream of the crossing position 18 towards the heat exchanger 27. Further, the gas flow area should be substantially the same or may even be decreased from an outlet of the heat exchanger 27 to an inlet of the downstream compressor (high pressure compressor) 11.

The heat exchanger 27 preferably comprises a plurality of exchangeable cooling modules. According to a first example, a plurality of such cooling modules is arranged as sectors, which together form an annular cooling structure. According to an alternative, or complement to the first example, a plurality of such cooling modules is arranged next to each others in an axial direction of the gas turbine engine.

According to a further alternative, the heat exchanger may comprise a plurality of heat exchanging element arranged in series in said loop 26.

According to the embodiment shown in figure 1, the heat exchanger 27 is positioned radially interior of the intermediate casing 5. More specifically, the bypass channel is divided and the heat exchanger 27 is positioned in a secondary by-pass channel. In other words, the heat exchanger 27 is positioned radially interior of the main bypass passage 7. This position of the heat exchanger is possible due to the fact that a downstream end of the low pressure compressor 10 is positioned on a substantially larger radial distance than an upstream end of the high pressure compressor 11, see figure 1.

The design of the second channels 17 in figure 4 should only be regarded as a schematic example showing the function of the invention. In case they are positioned in a flow channel (such as the bypass channel) , they are preferably aerodynamically shaped for causing as small disturbance as possible to the passing bypass flow.

The inventive component 15 is configured for a core engine, i.e an engine with or without a bypass channel.

In the application shown in figure 1, the component is arranged so that uncombusted air is conveyed via the second channels 17 during operation. Further, the component is arranged so that the second gas flow in the second channels flows from a radially outer position to a radially inner position in relation to the crossing point during operation. Further, the gas turbine comprises a channel loop 26 connecting said first channels 16 to said second channels 17 so that a gas flow from said first channels subsequently will flow through said second channels in gas turbine operation. Further, the gas turbine is configured to cool the gas flow in said loop 26 during operation.

A second embodiment of the component is indicated in figure 6. The first channels 116 are directed at an oblique angle with regard to the axial direction. Thus,

• the direction of the first channels 116 has a component in the radial direction while the second channels are directed substantially in the radial direction. In other words, the first channels 116 cross the second channels 117 in an angle substantially, different from 90°.

According to one example, the loop 26 connects two adjacent channels 16,17 in the circumferential direction. In other words, a gas flow in one of said first channels ^' 16 returns in a neighbouring second channel 17. According to an alternative, the loop may be arranged so that the gas flow in one of said first channels 16 may be returned in a second channel 17 at any circumferential distance from the first channel 16. According to a further alternative, the component may be configured so that the gas flow from a plurality of said first channels 16 is joined before it is returned in one or a plurality of said second channels 17. The invention is not in any way limited to the above described embodiments, instead a number of alternatives and modifications are possible without departing from the scope of the following claims.

For example, the heat exchanger may be located in a secondary gas flow channel (by-pass channel) or in engine configurations without a by-pass channel, on the exterior of the engine housing. Further, the design of the heat exchanger should only be regarded as an example .

Further, the direction of the first channels and the direction of the second channels in the crossing point may vary in different applications. Further, the direction of different first cannels and/or second channels may vary along the circumference of the component .

Further, according . to an alternative to the wall structure embodiment shown in figure 2-5, the first and second channels may be formed by tubes crossing each other. Further, the component 15 may be formed by joining a plurality of sub-components. For example, the first channels may me integrated in a first subcomponent and the second channels may be integrated in a second sub-component.

According to an alternative to the shown embodiments, the channel division is located upstreams (or downstreams) of the position where the radial extension of the first channels is increased.

Claims

1. An annular gas turbine housing component (15) comprising a plurality of circumferentially spaced first channels (16) for conveying a first gas flow in a first direction characterized in that the component comprises a plurality of circumferentially spaced second channels (17) for conveying a second gas flow in a second direction across the first direction in a crossing position (18) of the first and second channels .

2. A gas turbine housing component according to claim ^' 1, characterized in that a plurality of the^' first channels (16) alternate with a plurality of the second channels (17) in the circumferential direction.

3. A gas turbine housing component according to claim 1, characterized in that every other channel in the circumferential direction is a first channel (16) and every other channel in the circumferential direction is a second channel (17) at least along a portion of the circumference .

4. A gas turbine housing component according to any preceding claim, characterized in that the component comprises a wall structure (19) , which comprises a plurality of circumferentially spaced radial walls (20,120,220), which define said first channels (16) and second channels (17) .

5. A gas turbine housing component according to claim 4, characterized in that the wall structure (19) is configured so that each of said radial walls

(20,120,220) define one of said first channels (16) on one side and one of said second channels (17) on the other side in the circumferential direction at least along a portion of the circumference.

6. A gas turbine housing component according to any preceding claim, characterized in that at least a plurality of the first channels (16) are configured so that there is an increase (21) in their radial extension from a position (22) upstream the crossing position in a gas flow direction to the crossing position (18 ).

7. A gas turbine housing component according to any preceding claim, characterized in that said plurality of first channels (16) are configured so that a flow path area in said crossing position (18) is at least as large as in a position (22) upstream the crossing position in a gas flow direction.

8. A gas turbine housing component according to any preceding claim, characterized in that said plurality of first channels (16) are configured so that a flow path area in a position (22) upstream the crossing position in a gas flow direction is substantially the same as in said crossing position (18) .

9. A gas turbine housing component according to any preceding claim, characterized in that the component comprises a circumferentially substantially continuous gas flow channel (23) , which is divided into said plurality of circumferentially spaced first channels (16) .

10. A gas turbine housing component according to any one of claims 6-8 and claim 9, characterized in that an axial position of said channel division is about at the same axial position as the upstream position (22) .

11. A gas turbine housing component according to claim 10, characterized in that a center line (24) of the continuous gas flow channel (23) in a radial direction in the upstream position (22) is positioned on substantially the same radial position as a center line (25) of said plurality of circumferentially spaced first channels (16) in said crossing position (18) .

12. A gas turbine housing component according to any preceding claim, characterized^' in that the first channels (16) are configured for conveying the first gas flow substantially in an axial direction (2) of the component .

13. A gas turbine housing component according to any preceding claim, characterized in that the plurality of second channels (17) are configured for conveying the second gas flow substantially in a radial direction of the component .

14. A gas turbine housing component according to any preceding claim, characterized in that the component is configured for a channel loop (26) connecting said first channels (16) to said second channels (17) so that a gas flow from said first channels subsequently will flow through said second channels in gas turbine operation.

15. A gas turbine housing component according to claim

14, characterized in that heat exchanger (27) is positioned in said loop for cooling the gas flow.

16. A gas turbine housing component according to any preceding claim, characterized in that the first gas flow channel (23) defines a primary gas flow channel (7).

17. A gas turbine characterized in that it comprises a gas turbine housing component (15) according to any preceding claim.

18. A gas turbine according to claim 17, characterized in that the component (15) is arranged so that uncombusted air is conveyed via_ the second channels (17) during operation.

19. A gas turbine according to claim 17 or 18, characterized in that the component is arranged so that the second gas flow in the second channels (17) flows from a radially outer position to a radially inner position in relation to the crossing point (18) during operation.

20. A gas turbine according to any one of claims 17-19, characterized in that the gas turbine comprises a channel loop (26) connecting said first channels (16) to said second channels (17) so that a gas flow from said first channels subsequently will flow through said second channels in gas turbine operation.

21. A gas turbine according to claim 20, characterized in that the gas turbine is configured to cool the gas flow in said loop (26) during operation.

22. A gas turbine according to any one of claims 17-21, characterized in that it comprises at least two axially spaced compressor stages (10,11) and that the gas turbine housing component (15) is positioned between the compressor stages (10,11) so that a gas flow from a first compressor stage (10) will enter the first channels (16) and the gas flow from the second channels (17) will enter a second compressor stage (11) in operation.

23. A gas turbine according to claim 22 characterized in that a downstream end of the first compressor stage (10) is positioned on a substantially larger radial distance than an upstream end of the second compressor stage (11) .

24. An annular gas turbine housing component comprising a plurality of circumferentially spaced first channels for conveying a first gas flow in a first direction, wherein the component comprises a plurality of circumferentially spaced second channels for conveying a second gas flow in a second direction across the first direction in a crossing position of the first and second channels.

25. A gas turbine housing component according to claim 24, wherein a plurality of the first channels alternate with a plurality of the second channels in the circumferential direction.

26. A gas turbine housing component according to claim

24, wherein every other channel in the circumferential direction is a first channel and every other channel in the circumferential direction is a second channel at least along a portion of the circumference.

^'27. A gas turbine housing component according to any one of claims 24-26, wherein the component comprises a wall structure, which comprises a plurality of circumferentially spaced radial walls, which define said first channels and second channels.

28. A gas turbine housing component according to claim

27, wherein the wall structure is configured so that each of said radial walls define one of said first channels on one side and one of said second channels on the other side in the circumferential direction at least along a portion of the circumference.

29. A gas turbine housing component according to any one of claims 24-28, wherein at least a plurality of the first channels are configured so that there is an increase in their radial extension from a position upstream the crossing position in a gas flow direction to the crossing position.

30. A gas turbine housing component according to any one of claims 24-29, wherein said plurality of first channels are configured so that a flow path area in said crossing position is at least as large as in a position upstream the crossing position in a gas flow direction.

31. A gas turbine housing component according to any one of claims 24-30, wherein said plurality of first channels are configured so that a flow path area in a position upstream the crossing position in a gas flow direction is substantially the same as in said crossing position.

32. A gas turbine housing component according to any one of claims 24-31, wherein the component comprises a circumferentially substantially continuous gas flow channel, which is divided into said plurality of circumferentially spaced first channels.

33. A gas turbine housing component according to any one of claims 29-31 and claim 32, wherein an axial position of said channel division is about at the same axial position as the upstream position.

34. A gas turbine housing component according to claim 33, wherein a center line of the continuous gas flow channel in a radial direction in the upstream position is positioned on substantially the same radial position as a center line of said plurality of circumferentially spaced first channels in said crossing position.

35. A gas turbine housing component according to any one of claims 24-34, wherein the first channels are configured for conveying the first gas flow substantially in an axial direction of the component.

36. A gas turbine housing component according to any one of claims 24-35, wherein the plurality of second channels are configured for conveying the second gas flow substantially in a radial direction of the component .

37. A gas turbine housing component according to any one of claims 24-36, wherein the component is configured for a channel loop connecting said first channels to said second channels so that a gas flow from said first channels subsequently will flow through said second channels in gas turbine operation.

38. A gas turbine housing component according to claim 37, wherein heat exchanger is positioned in said loop for cooling the gas flow.

39. A gas turbine housing component according to any one of claims 24-38, wherein the first gas flow channel defines a primary gas flow channel.

40. A gas turbine wherein it comprises a gas turbine housing component according to any one of claims 24-39.

41. A gas turbine according to claim 40, wherein the component is arranged so that uncombusted air is conveyed via the second channels during operation.

42. A gas turbine according to claim 40 or 41, wherein the component is arranged so that the second gas flow in the second channels flows from a radially outer position to a radially inner position in relation to the crossing point during operation.

43. A gas turbine according to any one of claims 40-42, wherein the gas turbine comprises a channel loop connecting said first channels to said second channels so that a gas flow from said first channels subsequently will flow through said second channels in gas turbine operation.

44. A gas turbine according to claim 43, wherein the gas turbine is configured to cool the gas flow in said loop .during operation.

45. A gas turbine according to any one of claims 40-44, wherein it comprises at least two axially spaced compressor stages and that the gas turbine housing component is positioned between the compressor stages so that a gas flow from a first compressor stage will enter the first channels and the gas flow from the second channels will enter a second compressor stage in operation.

46. A gas turbine according to claim 45 wherein a downstream end of the first compressor stage is positioned on a substantially larger radial distance than an upstream end of the second compressor stage.