Patent Application
20200024993
The present disclosure concerns a gas turbine engine combustion arrangement and a gas turbine engine.
A known gas turbine engine combustion arrangement comprises an annular combustion chamber, an annular discharge nozzle and an array of turbine nozzle guide vanes. The annular combustion chamber comprises an outer annular wall and the outer annular wall of the annular combustion chamber has a downstream end. The array of turbine nozzle guide vanes have radially outer platforms and the array of turbine nozzle guide vanes is located downstream of the annular combustion chamber. The annular discharge nozzle is located between the downstream end of the outer annular wall of the annular combustion chamber and the array of turbine nozzle guide vanes. The annular discharge nozzle having an upstream end and a downstream end and the upstream end of the annular discharge nozzle has an annular axially extending slot and the downstream end of the outer annular wall of the annular combustion chamber locates in the annular axially extending slot in the upstream end of the annular discharge nozzle. The radially outer platforms of the turbine nozzle guide vanes having upstream ends and the upstream ends of the radially outer platforms of the turbine nozzle guide vanes are positioned adjacent the downstream end of the annular discharge nozzle. A support structure is connected to the annular discharge nozzle and the support structure has an annular radially extending slot, the radially outer platforms of the turbine nozzle guide vanes have radially extending flanges and the radially extending flanges of the turbine nozzle guide vanes locate in the annular radially extending slot in the support structure.
However, this gas turbine engine combustion arrangement has a problem in that the radially extending flanges of the turbine nozzle guide vanes may become disengaged from the radially extending slot in the support structure. This may result in the leakage of cooling air due to an increase in an axial gap between the upstream ends of the radially outer platforms of the turbine nozzle guide vanes and the downstream end of the annular discharge nozzle producing a cold spot or cold spots. Additionally, an experienced fitter and a laborious assembly technique are required to assemble, or disassemble, the support structure onto the turbine nozzle guide vanes to ensure that the support structure is not damaged. After assembly boroscope inspection of the arrangement is required to ensure that the flanges of the turbine nozzle guide vanes are located in the radially extending slot in the support structure. Additionally, the support structure and discharge nozzle may only be used once due to wear produced due to the relative movement between the flanges of the turbine nozzle guide vanes and the radially extending slot in the support structure and because the support structure comprises a bonded, brazed or welded, sheet metal construction which is bonded, brazed or welded to the discharge nozzle.
According to a first aspect there is provided a gas turbine engine combustion arrangement comprising an annular combustion chamber, an annular discharge nozzle, an array of turbine nozzle guide vanes and an annular casing, the annular combustion chamber comprising an outer annular wall, the outer annular wall of the annular combustion chamber having a downstream end, the array of turbine nozzle guide vanes having radially outer platforms, the array of turbine nozzle guide vanes being located downstream of the annular combustion chamber, the annular discharge nozzle being located between the downstream end of the outer annular wall of the annular combustion chamber and the array of turbine nozzle guide vanes, the annular casing surrounding and being spaced radially from the annular combustion chamber, the annular discharge nozzle and the array of turbine nozzle guide vanes, the annular discharge nozzle having an upstream end and a downstream end, the upstream end of the annular discharge nozzle having an annular axially extending slot, the downstream end of the outer annular wall of the annular combustion chamber locating in the annular axially extending slot in the upstream end of the annular discharge nozzle, the radially outer platforms of the turbine nozzle guide vanes having upstream ends, the upstream ends of the radially outer platforms of the turbine nozzle guide vanes being positioned adjacent the downstream end of the annular discharge nozzle and a support structure being connected to the annular discharge nozzle and the annular casing.
The downstream end of the annular discharge nozzle may have an annular axially extending slot and the upstream ends of the radially outer platforms of the turbine nozzle guide vanes locating in the annular axially extending slot in the downstream end of the annular discharge nozzle.
A wear resistant coating may be provided on the upstream ends of the radially outer platforms of the turbine nozzle guide vanes and/or a wear resistant coating may be provided on the surfaces of the annular axially extending slot in the downstream end of the annular discharge nozzle.
A seal may be provided between the downstream end of the annular discharge nozzle and the upstream ends of the radially outer platforms of the turbine nozzle guide vanes.
The upstream ends of the radially outer platforms of the turbine nozzle guide vanes may be positioned downstream of or abutting the downstream end of the annular discharge nozzle.
The seal may abut radially outer surfaces at the upstream ends of the radially outer platforms of the turbine nozzle guide vanes.
The seal may comprise a brush seal. The brush seal may comprise a plurality of bristles and the bristles being arranged at an angle between 15° and 75° to a radial direction.
The seal may comprise a bellows seal. The bellows seal may have an S-shape cross-section, a C-shape cross-section, a W-shape cross-section or an E-shape cross-section.
The seal may be located in an annular member and the annular member is U-shape in cross-section. The annular member may have an axially extending flange.
The axially extending flange of the annular member may be bonded, brazed or welded to the annular discharge nozzle.
The downstream end of the annular discharge nozzle may have an annular axially extending slot and the axially extending flange of the annular member locating in the annular axially extending slot in the downstream end of the annular discharge nozzle. The axially extending flange of the annular member may be bonded, brazed or welded to the annular discharge nozzle.
The annular member may be removably secured to the annular discharge nozzle. The annular member may have at least one L-shape member and the at least one L-shape member is removably secured to the annular discharge nozzle. The annular discharge nozzle may have at least one radially outwardly extending flange and the at least one L-shape member of the annular member is removably secured to the at least one radially outwardly extending flange of the annular discharge nozzle. The annular member may have a plurality of circumferentially spaced L-shape members and each L-shape member is removably secured to the annular discharge nozzle. The annular discharge nozzle may have at least one radially outwardly extending flanges and each L-shape member of the annular member is removably secured to the at least one radially outwardly extending flange of the annular discharge nozzle. The annular discharge nozzle may have a plurality of circumferentially spaced radially outwardly extending flanges and each L-shape member of the annular member is removably secured to a corresponding one of the radially outwardly extending flanges of the annular discharge nozzle.
The support structure allows relative radial movement between the annular casing and the annular discharge nozzle but provides axial stiffness between the annular casing and the annular discharge nozzle.
The support structure connecting the annular discharge nozzle and the annular casing may comprise at least one support member. The at least one support member may be integral with the annular discharge nozzle. The at least one support member may be bonded, brazed or welded to the annular discharge nozzle. The at least one support member may be J-shape in cross-section, U-shape in cross-section, S-shape in cross-section or W-shape in cross-section.
The at least one support member may comprise a first portion secured to the annular casing and a second portion secured to the annular discharge nozzle and the first and second portions are radially slidable relative to each other. The first portion of the at least one support member may be removably secured to the annular casing. The second portion of the at least one support member may be bonded, brazed or welded to the annular discharge nozzle. The first portion of the at least one support member may have a radial projection arranged to locate in a pocket, recess, in the second portion of the at least one support member. The second portion of the at least one support member may have a radial projection arranged to locate in a pocket, recess, in the first portion of the at least one support member.
The annular casing may comprise a casing portion, the casing portion has an axial end, the axial end of the casing portion has a radially outwardly extending flange, the radially outwardly extending flange has a radial face and a recess in the radial face, the first portion of the at least one support member locates in the recess in the radial face of the casing portion, the first portion of the at least one support member is removably secured to the radially outwardly extending flange and the second portion of the at least one support member is bonded, brazed or welded to the annular discharge nozzle
The at least one support member may be removably secured to the annular casing. The at least one support member may be integral with the annular discharge nozzle or may be bonded, brazed or welded to the annular discharge nozzle. The annular casing may have at least one radially inwardly extending flange and the at least one support member is removably secured to the at least one radially inwardly extending flange. The annular casing may comprise a casing portion, the casing portion has an axial end, the axial end of the casing portion has a radially outwardly extending flange, the radially outwardly extending flange has a radial face and a recess in the radial face, the at least one support member locates in the recess in the radial face of the casing portion and the support member is removably secured to the radially outwardly extending flange.
The support structure connecting the annular discharge nozzle and the annular casing may comprise a plurality of circumferentially spaced support members. Each of the plurality of support members may be integral with the annular discharge nozzle. Each of the plurality of support members may be bonded, brazed or welded to the annular discharge nozzle. Each of the plurality of support members may be J-shape in cross-section, U-shape in cross-section, S-shape in cross-section or W-shape in cross-section.
Each of the support members may comprise a first portion secured to the annular casing and a corresponding second portion secured to the annular discharge nozzle and the first and second portions are radially slidable relative to each other. The first portion of each of the support members may be removably secured to the annular casing. The second portion of each of the support members may be bonded, brazed or welded to the annular discharge nozzle. The first portion of each of the support members may have a radial projection arranged to locate in a pocket, recess, in the corresponding second portion of the support member. The second portion of each of the support members may have a radial projection arranged to locate in a pocket, recess, in the corresponding first portion of the support member.
Each of the support members may be removably secured to the annular casing. Each support member may be integral with the annular discharge nozzle or may be bonded, brazed or welded to the annular discharge nozzle. The annular casing may have at least one radially inwardly extending flange and each support member is removably secured to the at least one radially inwardly extending flange. The annular casing may have a plurality of circumferentially spaced radially inwardly extending flanges and each support member is removably secured to a corresponding one of the plurality of radially inwardly extending flanges. The annular casing may comprise a casing portion, the casing portion has an axial end, the axial end of the casing portion has a radially outwardly extending flange, the radially outwardly extending flange has a radial face and a plurality of circumferentially spaced recesses in the radial face, each support member locates in a corresponding one of the recesses in the radial face of the casing portion and each support member is removably secured to the radially outwardly extending flange.
The annular casing may comprise a casing portion, the casing portion has an axial end, the axial end of the casing portion has a radially outwardly extending flange, the radially outwardly extending flange has a radial face and a plurality of circumferentially spaced recesses in the radial face, the first portion of each support member locates in a corresponding recess in the radial face of the casing portion, the first portion of each support member is removably secured to the radially outwardly extending flange and the second portion of each support member is bonded, brazed or welded to the annular discharge nozzle.
The first portion of each support member may have circumferentially spaced stops and the corresponding second portion of each support member is located circumferentially between the circumferentially spaced stops. The second portion of each support member may have circumferentially spaced stops and the corresponding first portion of each support member is located circumferentially between the circumferentially spaced stops.
The present disclosure also provides a gas turbine engine having a gas turbine engine combustion arrangement. The gas turbine engine may be an aero gas turbine engine, a marine gas turbine engine, an industrial gas turbine engine or an automotive gas turbine engine. The aero gas turbine engine may be a turbofan gas turbine engine, a turbojet gas turbine engine, a turbo-shaft gas turbine engine or a turbo-propeller gas turbine engine.
The skilled person will appreciate that except where mutually exclusive, a feature described in relation to any one of the above aspects may be applied mutatis mutandis to any other aspect. Furthermore except where mutually exclusive any feature described herein may be applied to any aspect and/or combined with any other feature described herein.
Embodiments will now be described by way of example only, with reference to the Figures, in which:
With reference to
The gas turbine engine 10 works in the conventional manner so that air entering the intake 11 is compressed by the fan 12 to produce two air flows: a first air flow into the intermediate pressure compressor 13 and a second air flow which passes through the bypass duct 23 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow 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 16, intermediate 17 and low 18 pressure turbines drive respectively the high pressure compressor 14, the intermediate pressure compressor 13 and the fan 12, each by suitable interconnecting shaft 20, 21 and 22 respectively.
In particular, as shown in
The annular combustion chamber 30 and comprises a radially inner annular wall structure 40, a radially outer annular wall structure 42 and an upstream end wall structure 44. The radially inner annular wall structure 40 comprises a first annular wall 46 and a second annular wall 48. The radially outer annular wall structure 42 comprises a third annular wall 50 and a fourth annular wall 52. The second annular wall 48 is spaced radially from and is arranged radially around the first annular wall 46 and the first annular wall 46 supports the second annular wall 48. The fourth annular wall 52 is spaced radially from and is arranged radially within the third annular wall 50 and the third annular wall 50 supports the fourth annular wall 52. The upstream end wall structure 44 comprises an upstream end wall 41 and a plurality of heat shields 43. The heat shields 43 are spaced axially from and are arranged axially downstream of the upstream end wall 41 and the upstream end wall 41 supports the heat shields 43. The upstream end of the first annular wall 46 is secured to the upstream end wall 41 of the upstream end wall structure 44 and the upstream end of the third annular wall 50 is secured to the upstream end wall 41 of the upstream end wall structure 44. The upstream end wall structure 44 has a plurality of circumferentially spaced apertures 54 and each aperture 54 extends through the upstream end wall 41 and a respective one of the heat shield 43. The combustion chamber 15 also comprises a plurality of fuel injectors 56 and a plurality of seals 58. Each fuel injector 56 is arranged in a corresponding one of the apertures 54 in the upstream end wall structure 44 and each seal 58 is arranged in a corresponding one of the apertures 54 in the upstream end wall structure 44 and each seal 58 is arranged around, e.g. surrounds, the corresponding one of the fuel injectors 56. The fuel injectors 56 are arranged to supply fuel into the annular combustion chamber 15 during operation of the gas turbine engine 10. The second annular wall 48 comprises a plurality of rows of combustion chamber tiles 48A, 48B and 48C and the fourth annular wall 52 comprises a plurality of rows of combustion chamber tiles 52A, 52B and 52C. The combustion chamber tiles 48A, 48B and 48C are secured onto the first annular wall 46 by threaded studs, washers and nuts and the combustion chamber tiles 52A, 52B and 52C are secured onto the third annular wall 50 by threaded studs, washers and nuts. The heat shields 43 are secured onto the upstream end wall 41 by threaded studs, washers and nuts. The heat shields 43 are arranged circumferentially side by side in a row.
The array of turbine nozzle guide vanes 34 comprises a plurality of circumferentially arranged turbine nozzle guide vanes 34 and each turbine nozzle guide vane 34 comprises an aerofoil portion 59, a radially outer platform 60 and a radially inner platform 61. The aerofoil portion 59, the radially outer platform 60 and the radially inner platform 61 of each turbine nozzle guide vane 34 are generally integral, e.g. form a monolithic structure. The turbine nozzle guide vanes 34 may be formed by casting or additive layer manufacture. The radially outer platforms 60 of the array of turbine nozzle guide vanes 34 have one or more rows of cooling apertures 63 extending there-through. The cooling apertures 63 are angled to provide a film of coolant on the inner surfaces 65 of the radially outer platforms 60 of the turbine nozzle guide vanes 34.
The third annular wall 50 of the radially outer annular wall structure 42 of the annular combustion chamber 30 has a downstream end 51, the array of turbine nozzle guide vanes 34 is located downstream of the annular combustion chamber 30 and the annular discharge nozzle 32 is located between the downstream end 51 of the third annular wall 50 of the annular combustion chamber 30 and the array of turbine nozzle guide vanes 34. The annular casing 36 surrounds and is spaced radially from the annular combustion chamber 30, the annular discharge nozzle 32 and the array of turbine nozzle guide vanes 34, e.g. the annular casing 36 surrounds and is spaced radially from the third annular wall 50 of the annular combustion chamber 30, the annular discharge nozzle 32 and the array of turbine nozzle guide vanes 34.
The annular discharge nozzle 32 has an upstream end 62 and a downstream end 64, the upstream end 62 of the annular discharge nozzle 32 has an annular axially extending slot 66, the downstream end 51 of the third annular wall 50 of the annular combustion chamber 30 locates in the annular axially extending slot 66 in the upstream end 62 of the annular discharge nozzle 32. The radially outer platforms 60 of the turbine nozzle guide vanes 34 have upstream ends 68, the upstream ends 68 of the radially outer platforms 60 of the turbine nozzle guide vanes 34 are positioned adjacent to the downstream end 64 of the annular discharge nozzle 32 and a support structure 70 is connected to the annular discharge nozzle 32 and the annular casing 36. In this example the downstream end 51 of the third annular wall 50 has a flange 53 which is L-shape in cross-section, and which extends radially outwardly and in an axially downstream direction. The portion of the flange 53 extending in an axially downstream direction, at the downstream end 51 of the third annular wall 50, locates in the axially extending slot 66 in the upstream end 62 of the annular discharge nozzle 32. The downstream end 64 of the annular discharge nozzle 32 has an annular axially extending slot 72 and the upstream ends 68 of the radially outer platforms 60 of the turbine nozzle guide vanes 34 locate in the annular axially extending slot 72 in the downstream end 64 of the annular discharge nozzle 32.
A wear resistant coating may be provided on the surfaces of the upstream ends 68 of the radially outer platforms 60 of the turbine nozzle guide vanes 34 and/or a wear resistant coating may be provided on the surfaces of the annular axially extending slot 72 in the downstream end 64 of the annular discharge nozzle 32. A wear resistant coating may be provided on the surfaces of the downstream end 51 of the third annular wall 50 of the annular combustion chamber 30 and/or a wear resistant coating may be provided on the surfaces of the annular axially extending slot 66 in the upstream end 62 of the annular discharge nozzle 32. A wear resistant coating may be provided on the surfaces of the portion of the flange 53 extending in an axially downstream direction, at the downstream end 51 of the third annular wall 50, and/or a wear resistant coating may be provided on the surfaces of the annular axially extending slot 66 in the upstream end 62 of the annular discharge nozzle 32. Sufficient clearance is provided between the downstream end 51 of the third annular wall 50, or the flange 53, and the annular axially extending slot 66 in the upstream end 62 of the annular discharge nozzle 32 to allow for relative movement between the annular combustion chamber 30 and the annular discharge nozzle 32 during normal operation of the gas turbine engine 10. Sufficient clearance is provided between the upstream ends 68 of the radially outer platforms 60 of the turbine nozzle guide vanes 34 and the annular axially extending slot 72 in the downstream end 64 of the annular discharge nozzle 32 to allow for relative movement between the turbine nozzle guide vanes 34 and the annular discharge nozzle 32 during normal operation of the gas turbine engine 10.
The flange 53 at the downstream end 51 of the third annular wall 50 is provided with a row of circumferentially spaced cooling apertures 55 and the cooling apertures 55 allow coolant to flow over the inner surface of the annular discharge nozzle 32 to provide a film of coolant on the inner surface 33 of the annular discharge nozzle 32 and on the inner surfaces 65 of the radially outer platforms 60 of the stage of turbine nozzle guide vanes 34.
The support structure 70 as shown more clearly in
Alternatively, as shown in
It may be equally applicable to provide a single annular support which comprises a single annular first portion secured to the annular casing and a single annular second portion secured to the annular discharge nozzle and the first and second portions are radially slidable relative to each other. The second portion may have a C-shaped, or J-shaped, cross section. The first portion of the support may be removably secured to the annular casing. The second portion of the support member may be bonded, brazed or welded to the annular discharge nozzle. The first portion of the support member may have an annular radial projection arranged to locate in an annular pocket, annular recess, in the second portion of the support. Alternatively, the second portion of the support may have an annular radial projection arranged to locate in an annular pocket, annular recess, in the first portion of the support. In this arrangement it is necessary to provide one or more apertures through the annular support to allow a flow of coolant to the cooling apertures in the radially outer platforms of the turbine nozzle guide vanes. The sliding joint between the first and second portions of the single annular support allows for relative thermal expansion and contraction between the annular casing and the annular discharge nozzle. The first and second portions form an S-shaped support which allows relative radial movement between the annular casing and the annular discharge nozzle but provide axial stiffness between the annular casing and the annular discharge nozzle. Alternatively, it may be possible to have a single annular second portion which has a C-shaped, or J-shaped, cross section and a plurality of circumferentially spaced radial projections each one of which locates in a pocket, recess, in a corresponding one of a plurality of first portions of the support or a single annular second portion which has a C-shaped, or J-shaped, cross section and a plurality of circumferentially spaced pockets, recesses, each one of which receives a radial projection of a corresponding one of a plurality of first portions of the support.
Another support structure 170 as shown more clearly in
A further support structure 270 as shown in
An additional support structure 370 as shown in
The seal arrangement 472 comprises a brush seal 474. The brush seal 474 is annular and comprises a plurality of bristles and the bristles are arranged such that they extend longitudinally at an angle between 15° and 75° to a radial direction. The brush seal 474 is carried by an annular member 476. The annular member 476 is U-shape in cross-section and the limbs of the U-shape annular member 476 extend radially. The annular member 476 has an axially extending flange 478. The brush seal 474 is located between the limbs of the U-shape annular member 476. The axially extending flange 478 of the annular member 476 is secured to the annular discharge nozzle 432 and in particular is bonded, brazed or welded to the annular discharge nozzle 432.
The support structure 70 is the same as that described with reference to
The support structure 70 is the same as that described with reference to
The seal arrangement 672 comprises a bellows seal 674. The bellows seal 674 is annular and the bellows seal 674 has a C-shape cross-section. The bellows seal 674 is carried by an annular member 676. The annular member 676 is U-shape in cross-section and the limbs of the U-shape annular member 676 extend radially. The annular member 676 has an axially extending flange 678.
The bellows seal 674 is located between the limbs of the U-shape annular member 676. The axially extending flange 678 of the annular member 676 is secured to the annular discharge nozzle 632 and in particular is bonded, brazed or welded to the annular discharge nozzle 632.
The support structure 70 is the same as that described with reference to
In further alternative arrangements of the downstream end of the combustion chamber, the discharge nozzle and the array of turbine nozzle guide vanes, not shown, it may be possible to provide a seal arrangement with a W-shape cross-section or an E-shape cross-section or other suitable cross-section bellows seal in the U-shape annular member.
In variations of the arrangements of
In other variations of the arrangements of
An advantage of the present disclosure is that it eliminates the radially extending flanges of the turbine nozzle guide vanes and the radially extending slot in the support structure and therefore they cannot become disengaged. It reduces the possibility of leakage of cooling air between the upstream ends of the radially outer platforms of the turbine nozzle guide vanes and the downstream end of the annular discharge nozzle. It is easier to assemble, or disassemble, the support structure onto the annular casing than onto the turbine nozzle guide vanes and this reduces the risk of damage to the support structure. It is not necessary to inspect the flanges of the turbine nozzle guide vanes to ensure they are located in the radially extending slot in the support structure. Furthermore, the support structure and discharge nozzle may be used many times increasing the service life and reducing the life cycle costs.
Although the present disclosure has been described with reference to an annular combustion chamber comprising an annular outer wall of the double skin arrangement in which a plurality of tiles are secured to the annular outer wall it may be equally possible for the annular outer wall to be of the single skin arrangement in which the annular outer wall does not have any tiles. The annular outer wall of the annular combustion chamber may also comprises a plurality of circumferentially arranged combustion chamber segments each of which extends the full length of the combustion chamber and each of which may have a single skin arrangement or a double skin arrangement.
Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. By way of example such engines may have an alternative number of interconnecting shafts (e.g. two) and/or an alternative number of compressors and/or turbines. Further the engine may comprise a gearbox provided in the drive train from a turbine to a compressor and/or fan.
The present disclosure also provides a gas turbine engine having a gas turbine engine combustion arrangement. The gas turbine engine may be an aero gas turbine engine, a marine gas turbine engine, an industrial gas turbine engine or an automotive gas turbine engine. The aero gas turbine engine may be a turbofan gas turbine engine, a turbojet gas turbine engine, a turbo-shaft gas turbine engine or a turbo-propeller gas turbine engine.
It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.
| Number | Date | Country | Kind |
|---|---|---|---|
| 201811008200 | Mar 2018 | IN | national |
