JP7149156B2 - gas turbine combustor and gas turbine - Google Patents

Description

The present invention relates to gas turbine combustors and gas turbines.

A gas turbine is configured to generate combustion gas, which is a high-temperature working fluid, by combusting compressed air compressed by a compressor together with fuel in a combustor, and to drive a turbine with the combustion gas (Patent Document 1, etc.). reference).

JP-A-2002-89842

Turbine components that are exposed to high temperature combustion gases need to be cooled using cryogenic fluid, which has a lower temperature than the combustion gases, as cooling air so as to prevent cracking and erosion due to oxidation or thermal stress. In addition, a divided structure is often adopted for turbine parts for the purpose of easing stress concentration due to thermal deformation and facilitating maintenance, inspection, and replacement. For this reason, it is necessary to supply a high-pressure cryogenic fluid as seal air from the outside of the combustion gas flow path to the gap between the parts so that the combustion gas does not flow into the gap between the parts. However, since the low-temperature fluid supplied to the turbine as cooling air and sealing air is generally extracted from the compressor, excessively increasing the amount of extracted air reduces the flow rate of the combustion gas that drives the turbine, leading to a decrease in the efficiency of the entire gas turbine. . Therefore, seal plates span the gaps between the parts of the turbine. However, these seal plates do not completely seal the gaps between the parts of the turbine, and a minute cryogenic fluid leak occurs between the turbine parts, and this leaked air contributes to the cooling of the seal plates and turbine parts. do.

However, in gas turbines for power generation, where efficiency improvement is urgently required as environmental problems become more serious, combustion gas temperature (turbine inlet temperature) is increasing year by year. The heat load on the first stage vane of the located turbine is increasing. In particular, it is difficult to apply a cooling structure to the stator vane end wall of the first stage of the turbine, which has a gap between it and the combustor transition piece, because it is thin. Leaked air from the gap between the combustor transition piece can be used for film cooling of the stator vane end wall, but this area is a complex flow field that is strongly affected by secondary flows such as horseshoe vortices, making it stable. It is difficult to form a film cooling film.

SUMMARY OF THE INVENTION An object of the present invention is to provide a gas turbine combustor and a gas turbine capable of improving the cooling efficiency of the first stage stator blade end wall of the turbine by making use of the leaked air from the gap between the combustor transition piece and the turbine stator blade. to provide.

In order to achieve the above object, the present invention provides a gas turbine combustion engine comprising a combustor transition piece for supplying combustion gas to a turbine, and assembled with a gap between the combustor transition piece and the inlet of the turbine. wherein a plurality of shaped holes are provided in a flange connected to an outlet portion of the combustor transition piece, a plurality of cooling holes for cooling the combustor transition piece are provided in the combustor transition piece, and the plurality of A gas turbine combustor is provided in which shaped holes are inclined toward the turbine side from the outer peripheral side of the flange toward the inner peripheral side thereof at a larger angle than the cooling holes .
Further, in a gas turbine combustor provided with a combustor transition piece for supplying combustion gas to the turbine and assembled with a gap between the combustor transition piece and an inlet of the turbine, the exit of the combustor transition piece A plurality of shaped holes are provided in a flange connected to the part, and the plurality of shaped holes are provided in a plurality of rows in the turbine axial direction, and the positions of the shaped holes in rows adjacent to each other in the turbine axial direction are aligned in the turbine circumferential direction. To provide a gas turbine combustor that is misaligned.

According to the present invention, it is possible to improve the cooling efficiency of the first-stage stator vane end wall of the turbine by utilizing leaked air from the gap between the combustor transition piece and the turbine stator vane.

Schematic diagram of an example of a gas turbine to which a gas turbine combustor according to an embodiment of the present invention is applied 1 is a cross-sectional view of a connection portion between a gas turbine combustor and a turbine according to an embodiment of the present invention; 1 is a rear view of a combustor transition piece provided in a gas turbine combustor according to an embodiment of the present invention, viewed from the turbine side; Cross-sectional view taken along line IV-IV in Fig. 3 Arrow view by arrow V in FIG. Arrow view by arrow VI in Fig. 3 Cross-sectional view taken along line VII-VII in Fig. 5 A view showing the trailing edge of the combustor transition piece shown in FIG. 3 together with the stator blades of the turbine first stage.

Embodiments of the present invention will be described below with reference to the drawings.

-gas turbine-
FIG. 1 is a schematic diagram of an example of a gas turbine to which a gas turbine combustor according to an embodiment of the invention is applied. In this embodiment, the gas turbine combustor is abbreviated as a combustor. A gas turbine 100 shown in FIG. 1 includes a compressor 51 , a combustor 52 and a turbine 53 . The compressor 51 compresses the air A sucked through the intake section, generates high-pressure compressed air C, and supplies the compressed air C to the combustor 52 . The combustor 52 combusts the compressed air C compressed by the compressor 51 together with fuel to generate high-temperature combustion gas H and supply it to the turbine 53 . Turbine 53 is driven by combustion gas H generated in combustor 52 . The compressor 51 and the turbine 53 are coaxially connected, and the compressor 51 or the turbine 53 is connected with a load device (generator 54 in this embodiment). Part of the rotational power obtained by the turbine 53 is used as power for the compressor 51 and part is used as power for the generator 54 . The combustion gas H that drives the turbine 53 is discharged from the turbine 53 as exhaust gas E.

Although FIG. 1 illustrates a single-shaft gas turbine as an application target of the present invention, similar effects can be obtained when the present invention is applied to a combustor transition piece in other types of gas turbines. . For example, the present invention can be applied to a two-shaft gas turbine having a high-pressure turbine connected to a compressor and a low-pressure turbine separated from the high-pressure turbine and connected to load equipment, and similar effects can be obtained. .

-Combustor-
FIG. 2 is a cross-sectional view (cross-sectional view including the turbine rotating shaft) of the connection portion between the combustor and the turbine according to this embodiment. As shown in FIG. 2 , combustor 52 includes combustor transition piece 16 that supplies combustion gas H to turbine 53 . The combustor transition piece 16 is assembled with a gap S between it and the inlet of the turbine 53 . The turbine 53 has at least one turbine stage 57 including a row of stator blades 55 and a row of rotor blades 56 . In the same stage, the stator blade cascade 55 and the rotor blade cascade 56 are arranged in order from the upstream side in the turbine axial direction. In the present embodiment, a case is illustrated in which a plurality of turbine stages 57 are provided, and the first stage (first stage) stator blade row 55 and rotor blade row 56 are denoted by reference numerals. The stator blade cascade and rotor blade cascade also have substantially the same configuration.

The stationary blade cascade 55 includes a plurality of stationary blades arranged in the turbine circumferential direction (turbine rotation direction) and is formed in an annular shape, and is divided into a plurality of segments (FIG. 8) in the turbine circumferential direction. The moving blade cascade 56 also includes a plurality of moving blades arranged in the turbine circumferential direction. One segment of the stator blade cascade 55 has two stator blade end walls 17 on the inner peripheral side and the outer peripheral side and at least one blade portion 18 (see also FIG. 8). The stator blade end wall 17 is a thin plate-like member that defines the inner and outer peripheries of the combustion gas flow path (annular flow path through which the combustion gas H flows) in the stator blade cascade 55 . The blade portion 18 serves to rectify the combustion gas H, extends in the turbine radial direction, and connects the inner and outer peripheral stator blade end walls 17 . By engaging the stator blade end wall 17 on the outer peripheral side with the turbine casing 15 , the stator blade cascade 55 is fixed to the turbine casing 15 . The moving blade cascade 56 is configured by attaching a plurality of moving blades 12 to the outer peripheral portion of the disk 13 . The annular combustion gas flow path in the moving blade cascade 56 is defined by the shroud 14 on the outer peripheral side and by the outer peripheral surface of the disk 13 on the inner peripheral side. The shroud 14 is divided into a plurality of segments in the turbine circumferential direction, and each segment is fixed to the turbine casing 15 .

-Seal Mechanism-
As shown in FIG. 2 , the leading edges of the stator vane end walls 17 on the inner and outer peripheries of the first stage of the turbine and the trailing edges of the combustor transition pieces 16 face each other in the turbine axial direction with the above-described gap S interposed therebetween. A gap S between the trailing edge of the combustor transition piece 16 and the leading edge of the stationary blade end wall 17 is sealed with a sealing member 6 .

The combustor transition piece 16 has a flange 19 at a combustion gas outlet in the main body, and the flange 19 faces the stator blade end wall 17 . This flange 19 is also called a frame, and is connected (joined) to the outlet of the combustor transition piece main body and configured integrally with the main body. The flange 19 is formed in an arcuate rectangular shape, and when a plurality of combustor transition pieces 16 are arranged in the turbine circumferential direction, the outlet openings of each flange 19 are arranged in the circumferential direction to form an annular shape. Moreover, the flange 19 is provided with a protrusion 20 for holding the seal member 6 . The protrusions 20 extend from the inner and outer peripheral walls of the flange 19 in the turbine radial direction to the side opposite to the combustion gas flow path. That is, the projection 20 on the outer side in the turbine radial direction of the flange 19 protrudes outward in the turbine radial direction from the wall surface of the flange 19 facing the outer side in the turbine radial direction. On the contrary, the protrusion 20 on the inner side in the turbine radial direction of the flange 19 protrudes inward in the turbine radial direction from the wall surface of the flange 19 facing inward in the turbine radial direction.

As described above, the stator blade end wall 17 defines the inner and outer peripheries of the combustion gas flow path, and the high-temperature combustion gas H flowing through the combustion gas flow path and the low-temperature fluid L filling the space outside the combustion gas flow path. separated from These stator blade end walls 17 have seal grooves 21 on surfaces facing the combustor transition piece 16 . The seal groove 21 extends in the turbine circumferential direction and opens toward the combustor transition piece 16 .

The seal member 6 described above is a member formed by bending a plate material, for example, and includes a hook portion 7 and a seal plate portion 8 . The hook portion 7 is bent in a U shape, and is put on the projection portion 20 of the flange 19 of the combustor transition piece 16 to hold the projection portion 20 . The inner wall surface of the hook portion 7 is in contact with the protrusion 20 or has a slight gap between it and the protrusion 20 . The seal plate portion 8 is a portion extending in the turbine axial direction from the trailing edge of the hook portion 7 toward the stator blade end wall 17, is inserted into the seal groove 21 of the stator blade end wall 17, and is inserted into the combustor transition piece 16 ( It straddles the gap S between the flange 19 ) and the stator blade end wall 17 . The hook portion 7 of the seal member 6 seals the periphery of the protrusion 20 of the combustor transition piece 16 , and the seal plate portion 8 seals the gap S. However, these sealing members 6 do not completely seal the clearance S, and cause minute leaks of the low-temperature fluid L in the clearance S. As shown in FIG. This leak air L' is used to cool the stationary blade end wall 17 and the like of the first stage of the turbine.

－Cooling Mechanism of First Stage Stator Blade End Wall－
FIG. 3 is a rear view of the combustor transition piece provided in the combustor according to the present embodiment, viewed from the turbine side. 4 is a cross-sectional view taken along line IV-IV in FIG. 3, FIG. 5 is a view taken along arrow V in FIG. 3, FIG. 6 is a view taken along arrow VI in FIG. 3, and FIG. It is a cross-sectional view taken along line VII-VII in the middle. FIG. 8 is a diagram showing the trailing edge of the combustor transition piece shown in FIG. 3 together with the stator vane of the first stage of the turbine. 8, 5 and 6 are views viewed from the combustion gas flow path. In these figures, elements that have already been explained are denoted by the same reference numerals as in the previous drawings, and their explanations are omitted as appropriate.

The combustor transition piece 16 has a plurality of shaped holes 23 and 24 (FIGS. 5 to 8) and other cooling holes 25 (FIGS. 5 to 7) in addition to the flange 19 described above. The cooling holes 25 are for cooling the combustor transition piece 16 , and are widely distributed on the outer peripheral surface of the combustor transition piece 16 . The low-temperature fluid L flowing outside the combustor transition piece 16 flows into the combustor transition piece 16 through these cooling holes 25 . The purpose of providing the cooling holes 25 is to suppress the metal temperature of the combustor transition piece 16 by the low-temperature fluid L flowing inside. On the other hand, the shaped holes 23 and 24 are provided separately from the cooling holes 25 for the purpose of cooling not the combustor transition piece 16 but the stator vane end wall 17 of the first stage of the turbine. The cooling hole 25 has a circular orthogonal cross-section, whereas the shaped holes 23 and 24 have non-circular orthogonal cross-sections, and the exit openings of the shaped holes 23 and 24 are located at the leading edge (upstream side in the flow direction of the combustion gas). On the other hand, the trailing edge (downstream side) has a wide shape (trapezoidal shape in this example).

Unlike the cooling holes 25 , the shaped holes 23 and 24 are provided only in the flange 19 of the combustor transition piece 16 . These shaped holes 23 and 24 are arranged on the inner and outer peripheral walls of the flange 19 in the turbine radial direction corresponding to the inner and outer stator blade end walls 17 of the first stage of the turbine to be cooled. In this embodiment, the shaped holes 23 and 24 provided in the inner peripheral wall of the flange 19 in the turbine radial direction are illustrated, but the shaped holes 23 and 24 provided in the outer peripheral wall in the turbine radial direction have the same configuration. . As shown in FIG. 6, the shaped holes 23 and 24 are not provided in the side wall of the flange 19 facing the turbine circumferential direction. A region X in the figure is an R portion of the inner wall of the corner of the peripheral wall of the flange 19 (the corner formed by the inner peripheral wall in the radial direction of the turbine and the side wall).

The shaped holes 23 and 24 are arranged in rows in the circumferential direction of the turbine. That is, a plurality of rows (two rows in this embodiment) are formed by the shaped holes 23 and 24 in the turbine axial direction, and a further row of shaped holes 24 exists on the turbine 53 side of the row of shaped holes 23 . The positions of the shaped holes of rows adjacent to each other in the turbine axial direction are shifted in the turbine circumferential direction (that is, the positions of the shaped holes 23 and 24 are shifted in the turbine circumferential direction). In this embodiment, the pitches of the shaped holes 23 and 24 are the same, and the shaped holes 23 are shifted from the shaped holes 24 by half a pitch in the turbine circumferential direction. The rows of shaped holes 23 and 24 cover substantially the entire area in the turbine circumferential direction of the inner and outer wall surfaces in the turbine radial direction of the inner wall of the flange 19 .

As shown in FIG. 7 , the cooling holes 25 described above are inclined toward the turbine 53 from the outside to the inside (toward the combustion gas flow path) of the combustor transition piece 16 . The shaped holes 23 and 24 are inclined from the outer peripheral side of the flange 19 to the inner peripheral side toward the turbine 53 side at a larger angle than the cooling hole 25 . In other words, if the angle of inclination of the cooling hole 25 viewed from the turbine circumferential direction with respect to the normal line of the inner wall surface of the flange 19 is β (FIG. 7), the angle of inclination α of the shaped holes 23 and 24 similarly taken is less than the angle of inclination β. is also large (eg greater than 75°). In this way, the shaped holes 23 and 24 are elongated in the turbine axial direction, and at least a part of the shaped holes 23 and 24 overlaps the protrusion 20 in the turbine axial direction. In particular, the shaped holes 24 belonging to the row closest to the turbine 53 extend across the protrusions 20, the trailing edges of the shaped holes 24 are closer to the turbine 53 than the trailing edges (rear surfaces) of the protrusions 20, and the exits are It is approaching the stator vane end wall 17 of the first stage of the turbine.

The shaped hole 23, which overlaps less with the protrusion 20 in the turbine axial direction than the shaped hole 24, has a layout in which the inlet opens to the peripheral wall of the flange (thin plate portion excluding the protrusion 20). Therefore, the inclination angle α' of the shaped hole 23 may be slightly smaller than the inclination angle α of the shaped hole 24 for manufacturing reasons, but it is sufficiently larger than the inclination angle β of the cooling hole 25 . In this embodiment, when viewed from the turbine circumferential direction, the exit side of the shaped hole 23 is widened toward the turbine 53 as shown in FIG. The jetting angle of the jetted low-temperature fluid L'' is designed to suppress the difference.

-motion-
During operation of the gas turbine, part of the compressed air is extracted from a compressed air flow path (not shown) of the compressor 51 and supplied to various locations as seal air and cooling air. Compressed air discharged from the outlet of the compressor 51 is also supplied to the space around the combustor transition piece 16 as a high-pressure low-temperature fluid L, for example. When the low-temperature fluid L is supplied to the space around the combustor transition piece 16, the pressure difference between the inside and outside of the combustion gas flow path presses the seal member 6 toward the combustion gas flow path, thereby separating the combustor transition piece 16 from the first stage of the turbine. A gap S between the stationary blade end wall 17 is sealed. At the same time, the leaked air L' that cannot be completely sealed by the sealing member 6 flows through the gap S into the combustion gas flow path. Further, at a position immediately upstream of the stationary blade end wall 17 of the first stage of the turbine, the low-temperature fluid L'' flows into the combustion gas passage from the shaped holes 23, 24 so as to follow the flow of the combustion gas H. These shaped holes. The low-temperature fluid L″ ejected from 23 and 24 flows along the inner wall surface of the flange 19 in the combustion gas flow path, joins the leak air L′ leaking into the combustion gas flow path from the gap S, and joins the stator blades of the first stage of the turbine. A film cooling membrane covering the end wall 17 is formed.

-effect-
(1) According to experiments conducted by the inventors of the present application, with only the leaked air L′ from the gap S between the combustor transition piece 16 (flange 19) and the stationary blade end wall 17, the film cooling effect is , and did not extend to the leading edge of the airfoil 18 or the vicinity of the pressure surface. On the other hand, when the shaped holes 23 and 24 were formed in the flange 19 as in the present embodiment and the same test was conducted, it was found that the range where the film cooling effect extends is the leading edge of the blade portion 18 on the stationary blade end wall 17 of the first stage of the turbine and the front edge of the blade portion 18 . It was found that it spreads to the vicinity of the pressure surface. Since the shaped holes 23 and 24 are provided at positions immediately upstream of the stationary blade end wall 17, the low-temperature fluid L'' ejected from the shaped holes 23 and 24 and flowing along the wall surfaces near the surface of the stationary blade end wall 17. Compared to the case where only the leaked air L' from the gap S is present, even if the air flow rate is about the same, it is better to distribute and blow it out from the gap S and the shaped holes 23 and 24. However, the cooling performance improved over a wide range of the stationary blade end wall 17. As described above, according to the present embodiment, the leaked air L′ from the gap S is utilized to cool the first stage stationary blade end wall 17 of the turbine. Efficiency can be improved.

(2) The shaped holes 23 and 24 are inclined at a large inclination angle α, and the exit opening has a wider trailing edge than the leading edge, so that the cryogenic fluid L″ flows along the inner wall surface of the flange 19 . can be ejected, and a film cooling film can be effectively formed.

(3) By arranging the shaped holes 23 and 24 side by side in the turbine circumferential direction on the inner and outer circumferential sides of the flange 19 in the turbine radial direction, the shaped holes are positioned immediately upstream of the inner and outer stator blade end walls 17 in the turbine radial direction. Twenty-three, twenty-four rows may be provided. As a result, substantially the entire width of the stationary blade end wall 17 can be covered with the film cooling film, which is effective.

In this embodiment, the configuration in which the shaped holes 23 and 24 are provided in both the inner and outer circumferential walls of the flange 19 in the turbine radial direction has been described. However, in order to improve the film cooling effect compared to the conventional structure in which the shaped holes 23 and 24 do not exist at all, the shaped holes 23 and 24 on either the inner peripheral side or the outer peripheral side in the turbine radial direction are omitted. You can

(4) By arranging the shaped holes 23 and 24 in a zigzag pattern, it is possible to uniformize the distribution of the film cooling film in the circumferential direction of the turbine caused by the low-temperature fluid L″ ejected from the shaped holes 23 and 24. This allows the shaped holes to be uniformly distributed. It is possible to form a stable film cooling film as compared with the case where the film is formed in a single row.

However, as long as the basic effect (1) is obtained, either the row of the shaped holes 23 or the row of the shaped holes 24 may be omitted. Conversely, the number of rows of shaped holes may be three or more.

(5) The trailing edge of the shaped hole 24 is closer to the turbine than the trailing edge of the protrusion 20, and the shaped holes 23 and 24 are closer to the stator blade end wall 17 at the first stage of the turbine. As a result, a film cooling film can be formed on the stationary blade end wall 17 as widely as possible by the low temperature fluid L″ and the leaked air L′.

However, as long as the above effect (1) is obtained, even if the trailing edge of the shaped hole 24 is not closer to the turbine than the trailing edge of the protrusion 20, it is quiet enough to overlap the flange 19 in the axial direction of the turbine. It is sufficient if the shaped hole 24 is close to the blade end wall 17 . In addition to the shaped hole 24, there is no problem even if the trailing edge of the outlet opening of the shaped hole 23 is closer to the turbine 53 than the protrusion 20 is.

6 Seal member 16 Combustor transition piece 17 Stator blade end wall 19 Flange 20 Protrusion 23, 24 Shaped hole 25 Cooling hole 51 Compressor 52 Gas turbine combustion 53... Turbine 100... Gas turbine A... Air C... Compressed air H... Combustion gas S... Gap α, α'... Inclination angle of shaped hole β... Inclination angle of cooling hole

Claims (8)

A gas turbine combustor comprising a combustor transition piece for supplying combustion gas to a turbine and assembled with a gap between the combustor transition piece and an inlet of the turbine,
providing a plurality of shaped holes in a flange connected to the outlet of the combustor transition piece ;
a plurality of cooling holes for cooling the combustor transition piece are provided in the combustor transition piece;
A gas turbine combustor , wherein the plurality of shaped holes are inclined toward the turbine side from the outer peripheral side of the flange toward the inner peripheral side at an angle larger than that of the cooling holes . The gas turbine combustor of claim 1, wherein
the flange includes a protrusion that retains a seal member;
A gas turbine combustor, wherein the plurality of shaped holes are formed in the flange so that a positional range in a turbine axial direction overlaps with the projection. 2. The gas turbine combustor according to claim 1, wherein the exit openings of said plurality of shaped holes are shaped such that the trailing edges are wider than the leading edges. 2. The gas turbine combustor according to claim 1, wherein said plurality of shaped holes are arranged on at least one of an inner peripheral side and an outer peripheral side of said flange in a turbine radial direction. 2. The gas turbine combustor according to claim 1, wherein said plurality of shaped holes are arranged in a turbine circumferential direction. A gas turbine combustor comprising a combustor transition piece for supplying combustion gas to a turbine and assembled with a gap between the combustor transition piece and an inlet of the turbine,
providing a plurality of shaped holes in a flange connected to the outlet of the combustor transition piece;
A gas turbine combustor in which the plurality of shaped holes are provided in a plurality of rows in the turbine axial direction, and the positions of the shaped holes in adjacent rows in the turbine axial direction are shifted in the turbine circumferential direction. The gas turbine combustor of claim 6 ,
the flange includes a protrusion that retains a seal member;
A gas turbine combustor, wherein at least the trailing edges of the shaped holes belonging to the row closest to the turbine are closer to the turbine than the trailing edges of the protrusions. a compressor for compressing air;
A gas turbine combustor according to claim 1 or 6 ;
a turbine driven by combustion gases generated in the gas turbine combustor;
A gas turbine, comprising: a seal member for sealing a gap between the combustor transition piece and a stator blade end wall of the first stage of the turbine.

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