JPH1082527A - Gas turbine combustor - Google Patents

Abstract

PROBLEM TO BE SOLVED: To protect an amount of air not related to combustion from being increased, keep a characteristic of low NOx formation, improve a performance of a turbine and improve an air cooling performance and a strength of structure. SOLUTION: There are provided a cylindrical combustor liner 17, a pilot fuel nozzle 18 for supplying fuel into a combustion chamber, a pre-mixing duct 19 for supplying pre-mixing fuel of fuel and air and a tail cylinder 20 for feeding combustion gas to a turbine 13. A fuel blowing-out section of a pilot fuel nozzle 18 and an air flow passage for feeding combustion air into the pre-mixing duct are formed by an annular space 22 between the combustor liner 17 of the and a flow sleeve 21 arranged at its outer circumference. Combustion air is flowed within an air flow passage from a downstream side of the combustor liner 17 to its upstream side, thereby the combustion air is used as cooling air for the combustor liner 17. The outer circumferential surface of the liner 17 facing against the annular space is provided with rib-like fins 29 crossing with an axial direction of the combustor liner 17.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a gas turbine combustor applied to a thermal power plant, a combined cycle power plant and the like, and more particularly to a gas using combustion air as cooling air for a combustor liner or transition piece. The present invention relates to a technology for improving the cooling performance of a turbine combustor.

[0002]

2. Description of the Related Art In recent years, gas turbines applied to thermal power plants, combined cycle power plants, etc.
In order to achieve higher efficiency, operating conditions are increased in temperature and pressure, and the combustion gas temperature tends to increase accordingly.

As the combustion gas temperature increases, it is necessary to improve the structural strength and material strength and the cooling performance for preventing overheating from the viewpoint of extending the useful life of the gas turbine combustor and turbine components. In general, gas turbines often employ an air cooling method that uses compressed air from a gas turbine compressor to cool gas turbine combustors and turbine components.This method is used to prevent overheating of the gas turbine combustor. , It is necessary to supply a large amount of cooling air to the cooling function.

When the temperature of the combustion gas increases, the concentration of NOx contained in the exhaust gas tends to increase. Therefore, it is necessary to increase the ratio of fuel to fuel by supplying a large amount of combustion air as a preventive measure. As a technique for increasing the ratio of combustion air to fuel in order to reduce NOx in the exhaust gas, a premixed lean combustion method was developed in which combustion air and fuel are premixed and supplied to the combustor as lean fuel. Have been. According to this premixed lean burn system, the fuel and air supplied to the combustor liner are uniformly mixed, the local flame temperature is prevented from increasing, the combustion chamber temperature becomes uniform, and the low NO
x conversion is achieved.

[0005] In the gas turbine combustor of the premixed lean burn system, air supplied from the gas turbine compressor is also used as premixing air. Therefore,
Increasing the supply amount of the premixing air relatively reduces the amount of air that can be used as cooling air from the gas turbine compressor. Therefore, it is necessary to further improve the air cooling efficiency of the gas turbine combustor in order to increase the efficiency of the gas turbine while increasing both the cooling performance of the gas turbine and the amount of premixing air. Become.

FIG. 13 shows a cooling structure of the gas turbine combustor 1 by impingement cooling as an example of a high efficiency cooling means developed from such a viewpoint. In this gas turbine combustor 1, an outer peripheral side of a combustor liner 2 is covered with a sleeve 3, and an annular space 4 is formed therebetween. These sleeve 3 and combustor liner 2
A large number of air holes 5 and 6 are formed at different positions. Then, high-pressure air a flowing on the outer peripheral side of the sleeve 3 is ejected from the air hole 5 of the sleeve 3 toward the combustor liner 2, and the collision of the air a on the outer peripheral surface of the combustor liner 2 (impingement) is performed. Therefore, the combustor liner 2 can be cooled with high efficiency. The air “a” used for cooling is then supplied to the air holes 6 of the combustor liner 2.
From the combustion chamber 7 for combustion.

FIG. 14 shows, as another example, a cooling structure of a combustor by film cooling. In this technique, the combustor liner 2 has a multi-stage structure in the axial direction, an air hole 8 is formed in the step portion along the axial direction, and the combustor liner 2 is formed in the axial direction downstream of the air hole 8 in the combustor liner 2. A guide plate 9 is provided along the same. The air a flowing from the outer peripheral surface side of the combustor liner 2 through the air holes 8 is diffused and flows as a thin film along the inner peripheral surface of the combustor liner 2 by the guide plate 9, A film-like boundary layer is formed between the inner peripheral surface of the fuel cell 2 and the high-temperature combustion gas in the combustion chamber 7. The air a forming this boundary layer takes heat from the combustor liner 2 to perform a cooling action, and rises in temperature by itself to form a temperature gradient layer between the combustion gas and the combustor liner 2 to transfer heat. By reducing the rate, the temperature of the combustor liner 2 is reduced. This film cooling is often employed simultaneously with the impingement cooling shown in FIG.

[0008]

In order to obtain high cooling performance in the impingement cooling shown in FIG.
In order to increase the heat transfer coefficient, a high differential pressure is required between the outer peripheral space of the sleeve 3 and the inner peripheral space, that is, the annular space 4. However, this increase in the differential pressure leads to an increase in the combustor pressure loss, which leads to an increase in the discharge pressure of the gas turbine compressor or a decrease in the combustor outlet pressure, which adversely affects the gas turbine performance. there is a possibility.

On the other hand, in order to obtain high cooling performance in the film cooling shown in FIG. 14, it is necessary to increase the flow rate of cooling air along the inner peripheral surface of the combustor liner 2. However, since this cooling air is not involved in combustion,
When applied to the gas turbine combustor of the premixed lean combustion method described above, the equivalent ratio of the premixed fuel is increased, and a local increase in the flame temperature and a non-uniform temperature distribution occur, resulting in low NOx. May result in impairing the characteristics of the chemical conversion.

The present invention has been made in view of such circumstances, and it is possible to prevent an increase in the differential pressure and an increase in the amount of air which is not involved in combustion, thereby maintaining the characteristics of NOx reduction by the premixed lean burn system. It is an object of the present invention to provide a gas turbine combustor that can improve the turbine performance and also improve the air cooling performance and the structural strength.

[0011]

According to the first aspect of the present invention, there is provided a tubular combustor liner having an internal combustion chamber, and an upstream end portion of the combustor liner. A pilot fuel nozzle that supplies fuel into the combustion chamber, and a pilot fuel nozzle that is located downstream of the pilot fuel nozzle and is connected to a peripheral wall of the combustor liner to supply a premixed fuel of fuel and air. A premixed lean-burn gas turbine combustor comprising a mixing duct and a transition piece connected to a downstream end of the combustor liner and introducing combustion gas to the turbine side, wherein the pilot fuel nozzle has An air flow path for introducing combustion air into the fuel outlet and the premixing duct is formed by an annular space between the combustor liner and a flow sleeve provided on an outer peripheral side thereof,
The combustion air is used as cooling air for the combustor liner by flowing combustion air from the downstream side to the upstream side of the combustor liner in the air flow path. A gas turbine combustor characterized in that rib-shaped fins intersecting with the axial direction of a combustor liner are provided on the outer peripheral surface as turbulence promoting means for the cooling air.

According to the second aspect of the present invention, there is provided a cylindrical combustor liner having an internal combustion chamber, a pilot fuel nozzle provided at an upstream end of the combustor liner and supplying fuel into the combustion chamber, A premixing duct that is located downstream of the pilot fuel nozzle and is connected to the peripheral wall of the combustor liner and supplies a premixed fuel of fuel and air, and is connected to a downstream end of the combustor liner, A premixed lean burn type gas turbine combustor having a transition piece for introducing combustion gas to a turbine side, wherein air for introducing combustion air into a fuel blowing portion of the pilot fuel nozzle and the premixing duct is provided. A flow path is formed by an annular space between the combustor liner and the transition piece, and a flow sleeve and a transition piece outer cylinder respectively provided on the outer peripheral side thereof. The one used for cooling air of the transition piece and the combustor liner by flowing the transition piece from the downstream side toward the upstream side of the transition piece and the combustor liner, wherein the transition piece and the combustor liner facing the annular space are used. A gas turbine combustor characterized in that rib-shaped fins intersecting with the axial direction of a combustor liner are provided on the outer peripheral surface as turbulence promoting means for the cooling air.

According to a third aspect of the present invention, in the gas turbine combustor according to the first or second aspect, a space in the gas turbine casing surrounding the combustor liner and the transition piece is a flow of compressed air supplied from the gas turbine compressor. A plurality of air inlet holes communicating with the annular space from the flow path of the compressed air are formed at the downstream end of the flow sleeve of the combustor liner or the downstream end of the transition piece outer cylinder. There is provided a gas turbine combustor characterized in that:

According to a fourth aspect of the present invention, in the gas turbine combustor according to the third aspect, a plurality of air inlet holes are opened along the axial direction of the flow sleeve or the transition piece outer cylinder of the combustor liner, thereby forming an annular space. The additional inflow of the cooling air is performed in a forward flow direction along the air flow in the inside, and the number of the cooling air inflow holes, the opening area thereof, or the radial height of the annular space is determined by the flow sleeve and the tail. A gas turbine combustor is provided which is set based on a cooling air pressure, a flow rate, a flow velocity, or an air pressure balance in the annular space of a cylinder.

According to a fifth aspect of the present invention, in the gas turbine combustor according to any one of the second to fourth aspects, the annular space formed by the combustor liner and the flow sleeve is provided in the transition piece outer cylinder. Corresponding to the sum of the amount of air flowing in from the air inflow hole and the amount of air flowing in from the air inflow hole provided in the flow sleeve, the flow path cross-sectional area is smaller than the annular space on the outer peripheral side of the transition piece. With the setting expanded,
A gas turbine combustor characterized in that a frustoconical guide tube for directing air flowing through the annular space on the side of the transition piece to the outer peripheral surface of the transition piece is provided at an air inlet hole of the flow sleeve.

According to a sixth aspect of the present invention, in the gas turbine combustor according to any one of the first to fifth aspects, the turbulence promoting means provided on the outer peripheral surface of the combustor liner or the outer peripheral surface of the transition piece is provided. In addition to the rib-shaped fins, a rib-shaped fin as turbulence promoting means is provided on the inner peripheral surface of the flow sleeve or the inner peripheral surface of the transition piece outer cylinder in an arrangement facing the fins. A turbine combustor is provided.

According to a seventh aspect of the present invention, in the gas turbine combustor according to the sixth aspect, the fin provided on the outer peripheral surface of the combustor liner or the outer peripheral surface of the transition piece, and the inner peripheral surface of the flow sleeve or the transition piece outer cylinder. The relative position with respect to the fins provided on the inner peripheral surface side is a regular arrangement that is a plane target position that coincides in the combustor axial direction, or a staggered arrangement in which the pitch is shifted in the combustor axial direction. A gas turbine combustor is provided.

According to the invention of claim 8, in the gas turbine combustor according to any one of claims 1 to 7, the fins as turbulence promoting means are continuous or non-continuous in the circumferential direction of the combustor liner or transition piece. Provided as a continuous shape, or provided based on the distribution gradient of the axial heat transfer coefficient in the annular space portion, the shape, the size, the number of arrangement, the axial distance is changed, furthermore, the radial height of the annular space, A gas turbine combustor characterized by changing based on a distribution gradient of an axial heat transfer coefficient in the annular space.

According to a ninth aspect of the present invention, in the gas turbine combustor according to any one of the first to eighth aspects, the outer peripheral surface of the combustor liner or the outer peripheral surface of the transition piece and the inner peripheral surface or the tail of the flow sleeve are provided. A gas turbine combustor characterized in that a plurality of straightening plates for regulating air flow in an annular space along an axial direction of a combustor liner are provided on at least one of an inner peripheral surface side of a cylinder and an outer cylinder. I will provide a.

According to a tenth aspect of the present invention, in the gas turbine combustor according to the ninth aspect, the straightening plate is formed along the axial direction of the combustor liner, and is a straight axial rib or a combustor liner for moving cooling air straight. A gas turbine combustor characterized in that the gas turbine combustor is a curved axial rib that is spirally curved in an axial direction and generates a swirling flow in cooling air.

[0021]

DESCRIPTION OF THE PREFERRED EMBODIMENTS An embodiment of a gas turbine combustor according to the present invention will be described below with reference to FIGS.

First Embodiment (FIGS. 1 to 4) The gas turbine combustor of this embodiment is of a premixed lean burn type, and FIG. 1 is a cross-sectional view showing the entire configuration of the gas turbine combustor. FIG. 2 is a perspective view showing a part of the combustor liner of the gas turbine combustor shown in FIG. 1 in a sectional view, FIG. 3 is a sectional view showing a part of FIG. 2 in an enlarged manner, and FIG. It is sectional drawing which shows the transition piece part of the gas turbine combustor shown in FIG.

As shown in FIG. 1, the gas turbine combustor 11 is located between a gas turbine compressor 12 and a turbine 13 and is supported by a casing 14. Are located.

The gas turbine combustor 11 has a cylindrical combustor liner 17 having a combustion chamber 16 inside, and an upstream end of the combustor liner 17.
And a combustor liner 1 located downstream of the pilot fuel nozzle 18.
7 and a plurality of premixing ducts 19 for supplying a premixed fuel of fuel and air to the combustion chamber 16 and a downstream end of the combustor liner 17 for communicating the combustion gas to the turbine 13 side. And a transition piece 20 to be introduced into the main body.

The outer peripheral side of the combustor liner 17 is surrounded by a flow sleeve 21 as a liner outer cylinder, and an annular space 22 is formed between them. One end of the flow sleeve 21 is attached to the turbine casing 14 by the head plate 14.
a, and is fixed together with the pilot fuel nozzle 18. The head plate 14a is provided with a premixed fuel nozzle 23 for supplying the premixed fuel b in an arrangement surrounding the pilot fuel nozzle 18;
Faces the air inlet end 24 of the premixing duct 19.
The air inlet end 24 is disposed in the annular space 22.

On the other hand, the outer peripheral side of the transition piece 20 is covered by a transition piece outer cylinder 25, and an annular space 26 is formed therebetween. The tail tube outer tube 25 is connected to the flow sleeve 21 so that the two annular spaces 22 and 26 are in communication. That is, the transition piece 20 is fitted in a state of contact with the outer peripheral surface of the combustor liner 17, and the transition piece outer cylinder 25 is connected to the flow sleeve 21.
Are fitted in a state of contact with the inner peripheral surface. This allows
The two annular spaces 22, 26 communicate with each other.

In the present embodiment, an air inflow hole 27 is opened at the most downstream end position of the transition piece outer cylinder 25 in the axial direction of the combustor liner, and compressed air a sent from the gas turbine compressor 12 is Air flow path 28 in gas turbine casing
Through the air inflow holes 27 into the annular spaces 22 and 26. The air a that has flowed into the annular spaces 22 and 26 is applied to the outer peripheral surface of the transition piece 20 and the combustor liner 1.
After being cooled in contact with the outer peripheral surface of the nozzle 7, the air itself is preheated and supplied to the periphery of the pilot fuel nozzle 18 and the air inlet end 24 of the premixing duct 19. That is, the air a is used as cooling air after being used as cooling air in the annular spaces 22 and 26,
The annular spaces 22, 26 serve as cooling air passages.

As shown in FIGS. 2 and 3, on the outer peripheral surface of the combustor liner 17 facing the annular space 22, as a means for promoting cooling air turbulence, a rib-like shape intersecting the axial direction of the combustor liner 17 is provided. Fins 29 are provided. The fins 29 are formed integrally with the combustor liner 17 as ribs continuous along the circumferential direction of the combustor liner 17, and are arranged at a constant pitch along the axial direction. The cross section of each of the fins 29 has a rectangular shape, and the protruding height thereof is substantially constant.

As shown in FIG. 4, rib-shaped fins 30 are also provided on the outer peripheral surface of the transition piece 20 facing the annular space 26 along a direction intersecting the axial direction. The fins 30 are formed integrally with the transition piece 20 as ribs that are continuous along the circumferential direction of the transition piece 20, and are arranged at regular intervals along the axial direction. Also,
The cross-section of these fins 30 is also rectangular, and their protruding height is substantially constant.

Thus, the air a supplied from the gas turbine compressor 12 is annularly formed from the air flow path 28 in the gas turbine casing 14 through the air inflow hole 27 opened at the downstream end of the transition piece 20. It flows into the space 26. The inflowing air a cools the transition piece 20 while passing through the annular space 22 as cooling air, and then flows into the annular space 22 on the combustor liner 17 side to cool the combustor liner 17, and the air itself is preheated. And flows to the upstream side of the combustor liner 17. And, as described above, the pilot fuel nozzle 1
8 and flows into the combustion chamber 16 from the air blowing portion to be used for the combustion of the pilot fuel ejected from the pilot fuel nozzle 18, and the premix nozzle 19 disposed in the annular space 22. The fuel enters the premixing duct 19 from the air inlet end 24 and is premixed with the fuel b, becomes a premixed lean fuel, and is ejected from the outlet 19 a into the combustion chamber 16.

In the present embodiment, the outer surface of the transition piece 20 facing the annular spaces 26 and 22 and the outer peripheral surface of the combustor liner 17 intersects the axial direction of the combustor liner 17 as turbulence promoting means for cooling air. Since the rib-shaped fins 30 and 29 are provided, turbulence of the air flowing through the annular spaces 26 and 22 is promoted.

That is, the cooling air flows along the axial direction of the annular spaces 26 and 22, and the cooling air flows in layers along the outer peripheral surface of the transition piece 20 and the outer peripheral surface of the combustor liner 17. A boundary layer is formed on the outer peripheral surface of the combustor liner 17 and the outer peripheral surface of the combustor liner 17. The flow of the cooling air serving as the boundary layer is turbulent by the fins 30 and 29. That is, the boundary layer of the cooling air is
9 is temporarily separated by the tip of the fin 3
On the downstream side after exceeding 0 and 29, the state is reattached to the outer peripheral surface of the transition piece 20 and the outer peripheral surface of the combustor liner 17.
Since the boundary layer near the reattachment point becomes thinner and the heat transfer coefficient increases as the boundary layer becomes thinner, a plurality of fins 3 as turbulence promoting means are provided to the transition piece 20 and the combustor liner 17.
By arranging 0 and 29, the boundary layer is periodically peeled and reattached, so that the effect of promoting heat transfer becomes extremely large.

As described above, the cooling air is supplied to the air inlet 2
7, the annular space 26 of the transition piece 20 and the annular space 22 of the combustor liner 17 flow in this order.
And cools the combustor liner 17 to an appropriate temperature and acts to maintain their structural strength. Further, in the cooling structure of the present embodiment, the amount of the cooling air flowing into the combustion chamber 16 without directly participating in the combustion can be reduced, and the combustion air can be maximally used as the cooling air. The cooling performance of the fuel cell can be improved, and at the same time, effects such as premixing of fuel effective for reducing NOx can be achieved.

In this embodiment, the fins 29, 30 are provided on both the combustor liner 17 and the transition piece 20, respectively. However, the fins 29, 30 are provided only on either the combustor liner 17 or the transition piece 20. May be provided.

Second Embodiment (FIG. 5) FIG. 5 is an overall configuration diagram showing the configuration of a gas turbine combustor according to this embodiment.

The present embodiment differs from the first embodiment in that the air inlet holes 3 of the transition sleeve 25 and the flow sleeve 21 covering the combustor liner 17 are also provided.
2 is provided. The other configuration is almost the same as that of the first embodiment, and a duplicate description will be omitted.

Air flows into the annular space 22 from a plurality of locations along the axial direction of the combustor liner 17 through the air inflow hole 27 of the transition piece outer cylinder 25 and the air inflow hole 32 of the flow sleeve 21. Cooling air is additionally introduced in the forward flow direction of the air flowing from the annular space 26 on the side to the annular space 22 on the combustor liner 17 side. The number of these air inlet holes 27, 32, the opening area thereof, or the radial height of the annular spaces 26, 22 depends on the combustor liner 1
The cooling air pressure, the flow rate, the flow velocity, the air pressure balance in the annular spaces 26, 22 and the like of the cooling pipe 7 and the transition piece 20 are set to be appropriate.

The annular space 22 formed by the combustor liner 17 and the flow sleeve 21 is provided in the transition piece outer cylinder 2.
The annular space on the side of the transition piece 20 corresponds to the sum of the amount of air flowing from the air inlet hole 27 provided in the flow sleeve 21 and the amount of air flowing from the air inlet hole 32 provided in the flow sleeve 21. The flow path cross-sectional area is set to be larger than 26.

Further, a frustoconical guide tube 33 for directing the air flowing through the annular space 26 on the side of the transition piece 20 to the outer peripheral surface of the transition piece 20 is provided at the air inflow hole 32 of the flow sleeve 21. The frustoconical guide tube 33 provided at the junction has a function of smoothing the inflow of cooling air, and at the same time, reducing the turbulence caused by merging with the cooling air from the annular space 26 on the tail tube 20 side. To demonstrate.

In the present embodiment, the air a discharged from the gas turbine compressor 12 is annularly formed from two axial positions of the air inlet 27 of the transition piece outer cylinder 25 and the air inlet 32 of the flow sleeve 21. It flows into the spaces 26 and 22. Then, the air used for cooling the transition piece 20 joins the air flowing from the cooling air hole 32 of the flow sleeve 21 at the end of the transition piece 20 on the combustor liner 17 side. In this case, the frusto-conical guide tube 33 provided at the merging portion smoothes the inflow of the cooling air, and at the same time, reduces disturbance due to merging with the cooling air from the annular space 26 on the tail tube 20 side. .
Further, the air inlet hole 27 on the transition piece 20 side and the combustor liner 17
From the relationship with the air inlet holes 32 on the side, the pressure for obtaining the air flow rate in each air flow portion is properly balanced, and the pressure loss and cooling performance generated in the fins 30 and 29 for promoting turbulence are properly adjusted. Become.

Therefore, also in the gas turbine combustor of the present embodiment having the above cooling structure, the amount of cooling air flowing into the combustion chamber 16 without directly participating in combustion is reduced, and the combustion air is used as cooling air. In addition to maximizing the utilization, the cooling performance of the entire combustor can be improved, and premixing of fuel effective for lowering NOx can be achieved. Particularly, in the present embodiment, the cooling air is supplied from the air inlets 27 and 32 divided into two parts, so that a part of the air is supplied to the part of the combustor liner 17 which is located in the middle of the annular spaces 26 and 22. And the uncooled cooling air is supplied from this portion, thereby further improving the cooling performance.

Third Embodiment (FIGS. 6 and 7) FIG. 6 is a perspective view showing a part of a combustor liner 17 of a gas turbine combustor according to this embodiment, and FIG. 7 is a partial view of FIG. It is sectional drawing which expands and shows.

The gas turbine combustor of this embodiment is different from the above embodiments in that a plurality of gas turbine combustors are provided in the annular space 22 on the combustor liner 17 side so as to regulate the air flow along the axial direction of the combustor liner 17. Rectifying plate, that is, the axial rib 34
Is provided. Other configurations are almost the same as those of the above-described embodiments, and therefore, duplicate description will be omitted.

Axial rib 34 in the present embodiment
Are provided so as to protrude integrally with the outer peripheral surface of the combustor liner 17 and are arranged in a straight line along the axial direction.
A plurality of air conditioners are arranged at intervals in the circumferential direction so that the cooling air travels straight along the axial direction of the combustor liner 17.

According to the configuration of the present embodiment, the circumferential flow along the fins 29 for promoting turbulence of the combustor liner 17 and the circumferential direction when the annular space 22 is eccentric due to a processing assembly error. The non-uniform flow is suppressed or adjusted by the axial ribs 34, so that the flow of the cooling air can be made smooth and uniform along the axial direction of the combustor liner 17.

Further, according to the present embodiment in which the axial ribs 34 are provided integrally with the combustor liner 17, the air ribs also have the function of radiating fins at the same time as the function of rectifying the air flow, thereby promoting the cooling of the combustor liner 17. The effect is also achieved. Therefore, it is desirable that the dimension and shape such as the row pitch of the axial ribs 34 be appropriately determined in relation to the cooling performance required for the combustor liner 17 and the shape and cooling performance of the rib-like fins 29.

Although not shown, axial ribs may also be provided on the outer peripheral surface side of the transition piece in addition to the axial rib 34 of the present embodiment, and the axial rib 34 on the flow sleeve 21 side is omitted, and the transition rib on the transition piece side is omitted. It is also possible to provide only axial ribs.

Fourth Embodiment (FIGS. 8 and 9) FIG. 8 is a perspective view showing a partial cross section of a part of a combustor liner 17 in a gas turbine combustor of the present embodiment, and FIG. It is sectional drawing which expands and shows.

The gas turbine combustor according to the present embodiment is a modification of the third embodiment, in which an axial rib 35 is provided on the flow sleeve 21 side as a current plate. The other configuration is almost the same as that of the third embodiment, and the duplicate description will be omitted.

The axial ribs 35 of the present embodiment are also arranged in a straight line along the axial direction of the combustor liner 17, and a plurality of axial ribs are arranged at intervals in the circumferential direction so that the cooling air is supplied to the combustor liner 17. Is made to go straight along the axial direction.

According to such a configuration of the present embodiment,
The axial rib 35 suppresses or adjusts the circumferential flow along the turbulence-promoting fins 29 of the combustor liner 17 and the uneven flow along the circumferential direction when the annular space 22 is eccentric due to processing and assembly errors. Therefore, the flow of the cooling air can be made smooth and uniform along the axial direction of the combustor liner 17.

In the case of the present embodiment, when the flow sleeve 21 is finally assembled to the combustor liner 17, the axial ribs 35 are configured to partition the annular space 22. Can be simplified.

In addition to the axial rib 35 of the present embodiment, an axial rib (not shown) may also be provided on the outer peripheral surface side of the transition piece. The axial rib 35 on the flow sleeve 21 side is omitted, and only the axial rib on the transition piece side. It is also possible to provide.

Fifth Embodiment (FIG. 10) FIG. 10 is an enlarged sectional view showing a combustor liner 17 in a gas turbine combustor according to this embodiment.

In the gas turbine combustor of the present embodiment, the flow sleeve 21 is provided in addition to the rib-shaped fins 29 and 30 as turbulence promoting means provided on the outer peripheral surface of the combustor liner 17 (or transition piece 20). A rib-shaped fin 36 is provided on the inner peripheral surface side of the (or tail tube outer tube 25) as turbulence promoting means. Other configurations are almost the same as those of the above-described embodiments, and thus redundant description will be omitted.

In the illustrated example, the fins 36 provided on the flow sleeve 21 face the fins 29 provided on the combustor liner 17 and are regularly arranged at staggered positions shifted by half a pitch. Although not shown, the two fins 29 and 36 may be provided in a regular arrangement at the same pitch at the plane target positions coincident in the axial direction, or may be irregularly arranged at different pitches. Although not shown, the same applies to the case where fins are provided on the inner peripheral surface side of the transition piece outer cylinder.

According to this embodiment having such a structure, the fins 36 provided on the side of the flow sleeve 21 and the like function similarly to the fins 29 provided on the side of the combustor liner 17 and the like, and are perpendicular to the flow of the cooling air. Since a velocity distribution in the flow direction is generated, a function of further promoting the separation of the boundary layer generated by the fins 29 such as the combustor liner 17 and the reattachment action downstream thereof is provided. The effect of further improving the cooling effect of the combustor liner 17 and the like is achieved. In this embodiment, each fin 29, 36
Are opposed to each other during the entire assembly, so that there is also an advantage that the processing of the combustor liner 17 is not particularly complicated and the cooling effect can be improved.

Sixth Embodiment (FIG. 11) FIG. 11 is an enlarged sectional view showing a combustor liner 17 in a gas turbine combustor according to this embodiment.

In the gas turbine combustor of this embodiment, the cross-sectional area of the annular space 22 is gradually reduced along the flow direction of the cooling air (the axial direction of the combustor), and the combustor is similarly changed in the flow direction of the cooling air. The height of the fins 29 provided on the outer peripheral surface of the liner 17 is gradually reduced, and the fin pitch is gradually reduced. Other configurations are almost the same as those of the above embodiments.

According to the configuration of the present embodiment, it is easy to arbitrarily set the gradient and distribution form of the axial heat transfer coefficient distribution in the annular space 22, and the axial heat transfer gas temperature distribution inside the combustor can be easily determined. The resulting metal temperature distribution of the combustor liner 17 can be made uniform.

In connection with the above configuration, the fins as turbulence promoting means can be arbitrarily selected to be provided as a continuous or discontinuous shape in the circumferential direction of the combustor liner 17 or the transition piece 20. Based on the distribution gradient of the axial heat transfer coefficient in the spaces 22 and 26, the shape, size, number of arrays,
It is possible to provide with varying axial distance. Furthermore, the radial height of the annular space 22 can be variously changed based on the distribution gradient of the axial heat transfer coefficient in the annular space 22.

Seventh Embodiment (FIG. 12) FIG. 12 is a perspective view, partially in section, showing a combustor liner 17 in a gas turbine combustor according to this embodiment.

In the gas turbine combustor according to the present embodiment, the axial rib as the straightening plate is formed in a spiral curve along the axial direction of the combustor liner 17, and the curved axial rib 37 generates a swirling flow in the cooling air. It has been. Other configurations are almost the same as those of the above embodiments.

According to the configuration of this embodiment, the air flow becomes a swirling flow along the curved axial rib 37, and the cooling performance in the annular space 22 can be selectively changed.

In this embodiment, the case where the curved axial ribs 37 are provided on the inner peripheral surface side of the flow sleeve 21 is shown. However, this is a case where it is advantageous in that the processing of the combustor liner 17 is facilitated. For example, the curved axial rib 37 can be provided on the outer peripheral surface of the combustor liner 17.

It is needless to say that each of the above-described components can be provided on the transition piece side.

Other Embodiments In addition to the above embodiments, various modifications can be made in the present invention. For example, the cross-sectional shape of the fin is triangular, V-shaped,
Other various shapes, fins are arranged in an inclined manner in the annular space, and the like. In this manner, a configuration suitable for the type, size, and the like of the gas turbine combustor can be provided.

[0068]

As described above in detail, according to the present invention, the cooling performance by the cooling air on the outer periphery of the combustor liner or the transition piece is provided by the rib-shaped fins as turbulence promoting means arranged in the axial direction of the combustor. Therefore, it is possible to increase the premixed combustion air flow rate without increasing the combustor air pressure loss and the cooling water flow rate not directly involved in combustion. Therefore,
At the same time as improving the cooling performance, increasing the temperature of the combustor and increasing NOx
It is possible to easily establish the ultra-lean combustion condition for lowering the flame temperature, which is the main factor of the generation, and it is possible to effectively cope with a higher temperature and a lower NOx of the gas turbine combustor.

[Brief description of the drawings]

FIG. 1 is an overall configuration diagram showing a first embodiment of a gas turbine combustor according to the present invention.

FIG. 2 is a perspective view showing a combustor liner part according to the first embodiment in a partial cross section.

FIG. 3 is an enlarged sectional view showing a part of FIG. 2;

FIG. 4 is an enlarged sectional view showing a transition piece part in the first embodiment.

FIG. 5 is an overall configuration diagram showing a second embodiment of the gas turbine combustor according to the present invention.

FIG. 6 is a perspective view showing a partial cross section of a combustor liner in a third embodiment of the gas turbine combustor according to the present invention.

FIG. 7 is an enlarged sectional view showing a part of FIG. 6;

FIG. 8 is a perspective view showing a partial cross section of a combustor liner in a fourth embodiment of the gas turbine combustor according to the present invention.

FIG. 9 is an enlarged sectional view showing a part of FIG. 8;

FIG. 10 is an enlarged sectional view showing a combustor liner in a fifth embodiment of the gas turbine combustor according to the present invention.

FIG. 11 is an enlarged sectional view showing a combustor liner in a sixth embodiment of the gas turbine combustor according to the present invention.

FIG. 12 is a perspective view showing a partial cross section of a combustor liner in a gas turbine combustor according to a seventh embodiment of the present invention.

FIG. 13 is an enlarged sectional view showing a combustor liner for explaining a conventional example.

FIG. 14 is an enlarged cross-sectional view showing a combustor liner for explaining another conventional example.

[Explanation of symbols]

 11 Gas Turbine Combustor 12 Gas Turbine Compressor 13 Turbine 14 Casing 14a Head Plate 15 Turbine Shaft 16 Combustion Chamber 17 Combustor Liner 18 Pilot Fuel Nozzle 19 Premix Duct 20 Tail 21 Flow Sleeve 22 Annular Space 23 Premix Fuel Nozzle 24 Air inlet end 25 tail tube outer tube 26 annular space 27 air inlet hole 28 air flow passage 29 fin 30 fin 32 air inlet hole 33 guide tube 34 axial rib 35 axial rib 36 fin 37 axial rib

Claims (10)

[Claims] 1. A cylindrical combustor liner having an internal combustion chamber, a pilot fuel nozzle provided at an upstream end of the combustor liner and supplying fuel into the combustion chamber, A premixing duct positioned downstream and connected to the peripheral wall of the combustor liner for supplying a premixed fuel of fuel and air; and a premixing duct connected to a downstream end of the combustor liner to transfer combustion gas to a turbine side. A premixed lean burn type gas turbine combustor comprising a transition piece and a transition piece to be introduced to the pilot fuel nozzle, wherein an air flow path for introducing combustion air into a fuel outlet of the pilot fuel nozzle and the premixed duct is provided.
By forming an annular space between the combustor liner and the flow sleeve provided on the outer peripheral side thereof, by flowing combustion air from the downstream side of the combustor liner to the upstream side in the air flow path. In what is used as cooling air for the combustor liner, a rib-shaped fin intersecting with the axial direction of the combustor liner is provided on the outer peripheral surface of the combustor liner facing the annular space as a means for promoting turbulence of the cooling air. A gas turbine combustor characterized by being provided. 2. A tubular combustor liner having an internal combustion chamber, a pilot fuel nozzle provided at an upstream end of the combustor liner and supplying fuel into the combustion chamber, A premixing duct positioned downstream and connected to the peripheral wall of the combustor liner for supplying a premixed fuel of fuel and air; and a premixing duct connected to a downstream end of the combustor liner to transfer combustion gas to a turbine side. A premixed lean burn type gas turbine combustor comprising a transition piece and a transition piece to be introduced to the pilot fuel nozzle, wherein an air flow path for introducing combustion air into a fuel outlet of the pilot fuel nozzle and the premixed duct is provided.
The combustor liner and the transition piece are formed by an annular space between the flow sleeve and the transition piece outer cylinder provided on the outer peripheral side of the combustor liner and the transition piece, respectively. The cooling pipe is used as cooling air for the transition piece and the combustor liner by flowing from the downstream side to the upstream side of the liner. A gas turbine combustor, wherein rib-shaped fins intersecting with the axial direction of the combustor liner are provided as turbulence promoting means for air. 3. The gas turbine combustor according to claim 1, wherein a space in the gas turbine casing surrounding the combustor liner and the transition piece is a flow path for compressed air supplied from the gas turbine compressor. A plurality of air inflow holes communicating with the annular space from the flow path of the compressed air are formed at a downstream end of a flow sleeve of the combustor liner or a downstream end of the transition piece outer cylinder. And gas turbine combustor. 4. The gas turbine combustor according to claim 3, wherein the air inlet has a plurality of openings along the axial direction of the flow sleeve or the transition piece outer cylinder of the combustor liner, whereby the air flows in the annular space. So that the additional inflow of cooling air is performed in the forward direction along the number of the cooling air inflow holes,
The opening area or the radial height of the annular space is set based on the cooling air pressure, flow rate, flow velocity, or air pressure balance in the annular space of the flow sleeve and the transition piece. Gas turbine combustor. 5. The gas turbine combustor according to claim 2, wherein the annular space formed by the combustor liner and the flow sleeve flows from an air inflow hole provided in the transition piece outer cylinder. In accordance with the sum of the amount of air to be supplied and the amount of air flowing from the air inlet provided in the flow sleeve, the flow path cross-sectional area is set to be larger than the annular space on the outer peripheral side of the transition piece. A gas turbine combustor characterized in that a frustoconical guide tube for directing air flowing through the annular space on the side of the transition piece to the outer peripheral surface of the transition piece is provided at an air inflow hole portion of the flow sleeve. 6. The gas turbine combustor according to claim 1, wherein a rib-shaped fin is provided on an outer peripheral surface of the combustor liner or an outer peripheral surface of the transition piece. In addition, a gas turbine combustor characterized in that rib-shaped fins as turbulence promoting means are provided on the inner peripheral surface of the flow sleeve or the inner peripheral surface of the transition piece outer cylinder so as to face the fins. 7. The gas turbine combustor according to claim 6, wherein a fin provided on an outer peripheral surface of the combustor liner or an outer peripheral surface of the transition piece and an inner peripheral surface of the flow sleeve inner peripheral surface or the transition piece outer cylinder. The gas turbine combustor is characterized in that the relative position with respect to the fins provided in the gas turbine combustor is a regular arrangement that is a plane target position that coincides in the combustor axial direction, or a staggered arrangement with a pitch shifted in the combustor axial direction. 8. The gas turbine combustor according to claim 1, wherein the fins as turbulence promoting means are provided as a continuous or discontinuous shape in a circumferential direction of the combustor liner or the transition piece. Or, based on the distribution gradient of the heat transfer coefficient in the axial direction in the annular space portion, the shape, the size, the number of arrangement, the axial distance is provided to be changed, furthermore, the radial height of the annular space in the annular space portion A gas turbine combustor characterized by varying based on a distribution gradient of an axial heat transfer coefficient. 9. The gas turbine combustor according to claim 1, wherein an outer peripheral surface of the combustor liner or an outer peripheral surface of the transition piece, and an inner peripheral surface of the flow sleeve or a transition piece of the transition piece outer cylinder. A gas turbine combustor characterized in that a plurality of straightening plates for regulating an air flow in an annular space along an axial direction of a combustor liner are provided on at least one of a peripheral surface side. 10. The gas turbine combustor according to claim 9, wherein the straightening plate is formed along the axial direction of the combustor liner and is a straight axial rib for allowing cooling air to travel straight, or a spiral shape in the axial direction of the combustor liner. A gas turbine combustor characterized in that it is a curved axial rib that is curved to generate a swirling flow in cooling air.