GB2391297A

GB2391297A - Gas supply assembly

Info

Publication number: GB2391297A
Application number: GB0217097A
Authority: GB
Inventors: Anthony Pidcock; Desmond Close; Michael Paul Spooner; Paul Ashley Denman
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2002-07-24
Filing date: 2002-07-24
Publication date: 2004-02-04
Also published as: GB0217097D0

Abstract

A gas supply assembly for a combustor 15 of a gas turbine engine. The assembly has a cowl 26 arranged over the upstream end of the combustor, having a cross-sectional profile that is generally "V" shaped, and an apex 52 projecting in an upstream direction. The gas supply may be fed from a divergent member 40, and the cowl may include radially outer and inner tapering portions 48, 50. The tapering portions may each include an outwardly facing convex region 56 adjacent to a concave region 54 located at an outer edge of the inner and outer tapering portion, the concave faces merging into the convex regions. The radius of curvature of the convex or concave region may be greater than the radius of curvature of the apex region, and may be in the range of 5mm to 250mm, and the radius of curvature of the convex region may be in the range of 5mm to 250mm. Each of the tapered portions 48, 50 may have a profile that is generally of an elongated "S" shape. The gas feed member 40 may have an area ratio in the range of 1.1 to 2.2.The apex region of the cowl may have a radius of curvature in the range 2mm to 30mm.

Description

( 1 239 297

Gas Supply Assembly This invention relates to gas supply assemblies for 5 combustor turbine engines.

Gas turbine engines require gas exiting from the high pressure compressor to be distributed to annular channels radially inwardly and outwardly of the combustor. In order to effect such distribution of the gas, dump diffusers are lO provided. A disadvantage of such diffusers is that they incur a pressure loss of around 2%.

According to one aspect of this invention, there is provided a gas supply assembly for a combustor of a gas turbine engine, the assembly comprising a cowl arrangeable 15 over the upstream end region of the combustor, wherein the cowl has a cross-sectional profile which is generally of a V shape, having an apex region projecting, in use, in an upstream direction.

Preferably the cowl includes radially inner and outer DO tapering portions which preferably extend outwardly from the apex region.

Preferably, the assembly includes a gas feed member which may be divergent.

Preferably, the inner and outer tapering portions 25 include a concave region facing the gas feed member. The apex region may present a convex face to the gas feed member. The apex region of the cowl may have a radius of curvature in the range of substantially 2mm to 0 substantially 30mm, preferably in the range of substantially 4mm to substantially 15mm, more preferably in the range of substantially 6mm to substantially 12mm.

The concave regions of the inner and outer tapering portions are preferably arranged adjacent the apex region 3> and preferably merge therewith.

The inner and outer tapering portions may each include

an outwardly facing convex region provided adjacent the respective concave region of the inner and outer tapering portions. Conveniently, the convex regions are provided at outer edge regions of the inner and outer tapering 5 portions. The concave regions of the inner and outer tapering portions may merge into the respective convex regions. Each of the tapering portions may have a profile which is generally of an elongated S shape.

10 The gas feed member and the cowl may be annular in configuration. The radius of curvature of the, or each, concave and/or convex region of each tapering portion may be greater than the radius of curvature of the apex region.

Preferably, the radius of curvature of the, or each, concave and/or convex region of each tapering portion may be selected to allow a smooth flow of gas from the gas feed member around the cowl.

Preferably, the gas supply assembly includes a gas 90 feed member which may be divergent. The gas feed member may have an area ratio in the range of substantially 1.1 to substantially 2.2, preferably in the range of substantially 1.3 to substantially 1.9, and more preferably in the range of substantially 1.4 to substantially 1.7. The area ratio 25 of the gas feed member is defined herein as the ratio of the cross sectional area of the gas feed member at the trailing edge of the outlet guide vane of the adjacent compressor to the cross sectional area at the outlet of the gas feed member when the gas feed member is viewed parallel 30 to the principal axis of the engine.

Preferably, the radius of curvature of the concave region of each tapering portion is in the range of substantially 5mm to substantially 250mm, more preferably in the range of substantially 15mm to substantially 180mm 35 and most preferably substantially 30mm to 120mm.

Preferably, the radius of curvature of the convex region of

each tapering portion is in the range of substantially 5mm to substantially 250mm, more preferably in the range of substantially 15mm to substantially 180mm and most preferably in the range of substantially 30mm to 5 substantially 120mm.

An embodiment of the invention will now be described by way of example only, with a reference to accompanying drawing, in which: Fig. 1 is a sectional side view of the upper half of a 10 gas turbine engine; and Fig. is a sectional side view of the combustor of the gas turbine engine shown in Fig. 1 when viewed in a circumferential direction.

With reference to Fig. l, a ducted fan gas turbine 15 engine generally indicated at 10 has a principal axis X-X.

The engine 10 comprises, in axial flow series, an air intake ll, a propulsive fan 12, a compressor region 113 comprising an intermediate pressure compressor 13, and a high pressure compressor 14, combustion means llS 30 comprising a combustor 15, and a turbine region 116 comprising a high pressure turbine 16, an intermediate pressure turbine 17, and a low pressure turbine 18. An exhaust nozzle 19 is provided at the tail of the engine 10.

The gas turbine engine 10 works in the conventional 25 manner so that air entering the intake 11 is accelerated by the fan to produce two air flows: a first air flow into the intermediate pressure compressor 13 and a second air flow which provides propulsive thrust. The intermediate pressure compressor 13 compresses the air flow directed 30 into it before delivering that air to the high pressure compressor 14 where further compression take place.

The compressed air exhausted from the high pressure compressor 19 is directed into the combustor 15 where it is mixed with fuel and the mixture combusted. The resultant 5 hot combustion products then expand through, and thereby drive the high, intermediate and low pressure turbine 16,

17 and 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low pressure turbines 16, 17 and '8 respectively drive the high and intermediate pressure 5 compressors 14 and 13 and the fan 12 by suitable interconnecting shafts.

The intermediate and high pressure compressors 13, 14 each comprise a casing 20,22 which circumferentially surrounds and encloses axially alternating annular arrays 10 of rotor blades and stator vanes 34 (see Figs. 2 to 4), although only the blades 24 of the intermediate pressure compressor 13 and the blades 26 of the high pressure compressor 14 can be seen in Fig. 1.

Referring to Fig. 2, the combustor 15 is comprises an 15 annular combustion chamber 20 having radially inner and out wall structures 21 and 22 respectively. The combustion chamber 20 is secured to an outer engine casing 23 by a plurality of pins 24 (only one of which is shown). An annular cowl 26 is provided at the upstream end of tee 20 combustion chamber 20. A plurality of apertures 26A are defined in the cowl 26 and are circumferentially spaced around the cowl 26. Fuel is directed into the chamber 0 through a number of injector nozzles 25 (only one of which is shown and is represented in broken lines) located at the 25 upstream end of the combustion chamber 20. The fuel injector nozzles 25 are circumferentially spaced around the engine 10 and serve to spray fuel into air derived from the high pressure compressor 14 through the respective apertures 26A. The resultant fuel/air mixture is then 30 combusted within the chamber 20.

The inner and out wall structures 21 and 22 are Or generally the same construction and comprise an outer wall 27 and an inner wall 28. The inner wall 28 is made up of a plurality of discrete wall elements in the form of tiles 35 29, which are all of the same general rectangular configuration and are positioned adjacent each other. The

circumferentially extending edges 30, 31 of adjacent tiles overlap each other. Each tile 29 is provided with threaded studs 32 which project through apertures in the outer wall 27. Nuts 39 are screwed onto the threaded studs 32 and 5 tightened against the outer wall 27, thereby securing the tiles 29 in place.

A feed member in the form of an annular feed nozzle 40 is provided by inner and outer annular wall sections 41, 42, and extends from the high pressure compressor 14 to 10 direct air into the combustor 50. Fig. 2 shows a stator vane which is the furthest downstream of the stator vanes 34 and is referred to as an outlet guide vane 34A. The annular nozzle 40 diverges from the outlet guide vane 34A.

An inner annular space 44 is provided radially 15 inwardly of the inner wall 21 of the combustor 15 and is defined between the inner wall 21 and an inner engine casing 45. Similarly, an outer annular space 46 is provided radially outwardly of the outer wall 22 of the combustor 15 and is defined between the outer wall 22 and 90 the outer engine casing 23.

Air from the high pressure compressor 14 is directed by the annular nozzle 40 around the cowl 26 to the inner and outer annular spaces 44, 46, as indicated by the arrows A and B. The passage of the air from the annular nozzle 40 35 to the inner and outer annular spaces 44, 46 is assisted by the shape of the annular cowl 26. The cowl 26 has a cross-

sectional profile which, when viewed in a circumferential direction, as shown in Fig. 2, is generally of a V shape and comprises annular inner and outer tapering portions 48, 30 50 and an annular central portion or apex region 52. The apex region 52 is generally arranged in line with the centre of the nozzle 40, as, represented by broken line Y-Y in Fig. 2. The apex region 52 presents a convex face to the nozzle 40 and represents the null point for the flow of 35 air onto the cowl 26. The convex face of the apex region 52 is of a bluff configuration to prevent the null point of

( 6 the flow of air shifting beyond the apex region 52 during operation of the engine. The radius of curvature of the apex region 52 in the preferred embodiment is in the range of substantially 6mm to substantially 12mm.

5 The inner and outer tapering portions 48, 50 have a cross-sectional profile as shown in Fig. 2, which is of an elongated S configuration. Each of the inner and outer tapering portions 48, 50 comprises a concave region 54 facing towards the nozzle 40 which serve respectively to 10 direct air in radially inner and outer directions from the nozzle 40. The concave regions 54 merge into the apex region 52. Each of the inner and outer tapering portions 48, 50 also includes a convex edge region 56 provided adjacent the respective concave region 54, and merging IS therewith. The convex edge regions 56 direct air into the inner and outer annular spaces 44, 46. The convex and concave regions 54, 56 have radii of curvature which are greater than the radius of curvature of the apex 52. The radius of curvature of each of the concave and convex 20 regions 54, 56 in the preferred embodiment is in the range of substantially 30mm to substantially 120mm.

The nozzle 40 of the preferred embodiment has an area ratio in the region of substantially 1.4 to substantially 1.7. The area ratio of the nozzle 40 is determined by the 25 ratio of the cross-sectional area of the nozzle 40 at the outlet 60 designated by the broken line to the cross-

sectional area of the nozzle 40 at the position 62 designated by the broken line, which corresponds to the trailing edge of the outlet guide vane 34A. The aforesaid 30 cross-sections are those seen when the nozzle 40 is viewed generally paralled to the axis X-X of the engine 10.

While not wishing to be limited to any particular theory, it is believed that the construction as described above allows efficient splitting of air to the inner and 35 outer annular spaces 44, 46 by reason that the protruding apex 52 exerts a back pressure on the air exiting the

nozzle 40. This, it is believed, reduces the pressure difference between the air flowing along the central region of the annular nozzle 40 and the air flowing along the inner and outer walls 41, 42 of the nozzle 40.

s There is thus described a gas supply assembly for a combustor of a gas turbine engine which allows efficient supply of air to the inner and outer annular spaces surrounding the combustor of the gas turbine engine.

Various modifications can be made without departing tO from the scope of the invention, for example, the central portion 52 could present a planer, rather than a convex face to the annular nozzle 40.

Whilst endeavouring in the foregoing specification to

draw attention to those features of the invention believed IS to be of particular importance it should be understood that the Applicant claims protection in respect of any patentable feature or combination of features hereinbefore referred to and/or shown in the drawings whether or not particular emphasis has been placed thereon.

SO

Claims

1. A gas supply assembly for a combustor (15) of a gas turbine engine (10) , the assembly comprising a cowl (26) 5 arrangeable over the upstream end of the combustor (15) characterised in that the cowl (26) has a crosssectional profile which is generally of a V shape, having an apex region (52) projecting, in use, in an upstream direction.

2. A gas supply assembly according to claim l 10 characterised by a divergent gas feed member (40).

3. A gas supply assembly according to claim 2 characterised in that the cowl (26) includes radially inner and outer tapering portions (48, 50) which can extend from the apex region (52) to the upstream end region of the 5 compressor (15).

4. A gas supply assembly according to claim 3 characterised in that the inner and outer tapering portions (48,50) include a concave (54) region facing the gas feed member (40), the apex region (52) presenting a convex face 90 to the gas feed member (40).

5. A gas supply assembly according to claim 4 characterised in that the concave region (54) of the inner and outer tapering portions (48, 50) are arranged adjacent the apex region (52) and merge therewith.

95

6. A gas supply assembly according to claims 3, 4 or 5 characterised in that the inner and outer tapering portions (48,50) each include an outwardly facing convex region (56) provided adjacent the respective concave regions (54) of the radially inner and outer tapering portions (48,50), the 30 convex regions (56) being provided at outer edge regions of the inner and outer tapering portions and said concave faces (56) of the inner and outer tapering portions (48,50) merging into the respective convex regions (54).

7. A gas supply assembly according to claim 6 35 characterised in that the radius of curvature of the, or each, concave and/or convex region (54, 56) of each tapering

portion (48,50) in said cross-sectional profile is greater than the radius of curvature of the apex region (52).

8. A gas supply assembly according to claim 7 characterised in that the radius of curvature of the 5 concave region of each tapering portion is in the range of substantially 5mm to substantially 250mm.

9. A gas supply assembly according to claim 8 characterised in that the radius of curvature of the concave region (54) of each tapering portion (48,50) is in 10 the range of substantially 15mm to substantially 180mm.

10. A gas supply assembly according to claim 8 characterised in that the radius of curvature of the concave region (54) of each tapering portion (48,50) is in the range of substantially 30mm to substantially 120mm.

IS

11. A gas supply assembly accounting to any of claims 7 to 10 characterised in that the radius of curvature of the convex region (56) of each tapering portion (48,50) is in the range of substantially 5mm to substantially 250mm.

12. A gas supply assembly according to claim 11 90 characterised in that the radius of curvature of the convex region (56) of each tapering portion (48,50) is in the range of substantially 15mm to substantially 180mm.

13. A gas supply assembly according to claim 12 characterised in that the radius of curvature of the convex 2s region (56) of each tapering portion (48,50) is in the range of substantially 30mm to substantially 120mm.

14. A gas supply assembly according to any of claims 3 to 13 characterised in that each of the tapering portions (48,50) has a profile which is generally of an elongated S 30 shape.

15. A gas supply assembly according to any of claims 2 to 14 characterised in that the gas feed member (40) has an area ratio in the range of substantially 1.1 to substantially 2.2.

35

16. A gas supply assembly according to claim 15 characterised in that the gas feed member (40) has an area

ratio in the range of substantially 1.3 to substantially 1. 9.

17. A gas supply assembly according to claim 16 characterized in that the gas feed member has an area ratio S of substantially in the range of substantially 1.4 to substantially 1.7.

18. A combustor arrangement for a gas turbine engine (10), the arrangement comprising a combustor (15) and characterized by a gas feed assembly as claimed in any 10 preceding claim.

19. A gas supply assembly according to any preceding claim characterized in that the apex region has a radius of curvature in the range of substantially 2mm to substantially 30mm.

IS

20. A gas supply assembly according to claim 18 characterized in that the apex region has a radius of curvature in the range of substantially 4mm Go substantially 15mm.

21. A gas supply assembly according to claim 9 20 characterized in that the apex region has a radius or curvature in the range of substantially 6mm -o substantially 12mm.

22. A gas turbine engine (10) incorporating a combing arrangement as claimed in claim 18.

25

23. A gas supply assembly substantially as herein described with reference to Fig. 2 of the drawings.

24. A combustor arrangement substantially as here-e described with reference to Fig. 2 of the drawings.

25. A gas turbine engine (10) substantially as here-e 30 described with reference to Fig. 1 of the drawings.

26. Any novel subject matter or combination including novel subject matter disclosed herein, whether or not within the scope of or relating to the same invention as any of the preceding claims.