JPH08135968A - Gas turbine combustor - Google Patents

Abstract

PURPOSE: To reduce the flow rate of cooling air remarkably, contrive the improvement of cooling performance, permit the high temperature of a combustor and the increase of flow rate of premixed combustion air due to reduction of cooling air, obtain superdilute combustion condition by restraining the increase of flame temperature and resolve problems in both of increace of temperature of gas turbine combustor as well as the reduction of NOx. CONSTITUTION: A multitude of air holes are bored on an outer wall cylinder 20 along the axial direction of a combustor so as to permit the impingement cooling of an inner wall cylinder 19. A multitude of air holes, whose arrangement are different from the air holes on the outer wall cylinder, are bored to permit the film cooling of the inner wall cylinder 19. The mounting part of a supporting structural body, supporting the outer peripheral side of a combustion liner 13, is provided with the impinge cooling unit of the inner wall cylinder. A seal unit between a tail cylinder 16 at the downstream side end of the combustion liner 13 is provided with a slot type film cooling unit.

Description

Detailed Description of the Invention

[0001]

The present invention relates to a gas turbine plant,
The present invention relates to a gas turbine combustor applied to a combined plant or the like, and more particularly to a gas turbine combustor having a cooling performance effective for raising the temperature of combustion gas in a case where a fuel and air are mixed and burned.

[0002]

2. Description of the Related Art In recent years, gas turbines used in gas turbine plants, combined plants, and the like have a tendency to increase in combustion gas temperature due to operating conditions of high temperature and high pressure for high efficiency. Since the combustion gas temperature lowers the metal temperature of the combustor liner whose flow rate ratio to the combustion gas flow rate is determined, how to efficiently use the cooling air is a problem.

Efficiently lowering the metal temperature of the combustor liner is indispensable for maintaining the structural strength and material strength that affect the service life of the combustor, or for maintaining the combustor performance that enables highly efficient combustion. Since it is a technical issue, various combustor liner metal cooling methods have been developed,
Some have already been put to practical use.

FIG. 7 shows a cooling technique for a combustor liner which is applied to a conventional gas turbine combustor intended to raise the temperature of combustion gas. This technique applies impingement cooling and obtains a high heat transfer coefficient by utilizing a jet of cooling air.

That is, the combustor liner 1 is formed in a double cylinder shape, and the outer peripheral wall 2a and the inner peripheral wall 2b of the combustor liner 1 are made to have different positions from each other so that a large number of small air holes 3a, 3b are formed. Has been drilled. The cooling air a is a combustion chamber 5 inside the combustor liner 1 from the cooling / combustion air passage 4 side (low temperature side) formed in the outer peripheral space of the combustor liner 1 through the air holes 3a and 3b. When flowing into the combustion gas side (high temperature side), since the inner and outer air holes 3a and 3b are located at different positions, the cooling air a collides with the inner peripheral wall 2b, and a high heat transfer coefficient is obtained by the jet flow in that case. It has become.

FIG. 8 also illustrates another combustor liner cooling technique. In this technique, relatively low-temperature cooling air a is blown into the boundary layer of the combustion gas from the air holes 3c of the porous outer peripheral wall 2a and the annular gap 3d on the inner peripheral side thereof to perform so-called film cooling. It has become.

FIG. 9 shows still another example in which a turbulent flow promoting rib 6 is provided on the peripheral wall 2 of the combustor liner 1 to form a complicated air passage 4. In this example, the cooling air a is made turbulent to improve the efficiency of the heat transfer coefficient and improve the heat transfer characteristics.

Conventionally, in addition to the combustor made of a high heat resistant alloy which employs the above-described various cooling methods, a combustor which employs a non-cooling method using a heat insulating material for high heat resistance is also known. . The former gas turbine combustor made of high heat resistant alloy is used relatively often especially in the case of multi-can type with a large cooling area, and the latter is a single-can type or a gas turbine main body casing to which high temperature heat insulating material is easily applied. It is used in a large separately installed combustor.

In the former combustor, since various cooling methods are combined and employed, there are problems that the combustor has a large number of constituent parts, the shape is complicated, and the manufacturing cost is high. Also, the latter combustor has substantially the same problem, and further, due to the limitation of the heat resistant use range of the heat insulating material, it is not easy to raise the temperature of the combustion gas, and at the same time, there is a problem that a complicated assembly structure is formed by combining different materials. there were.

On the other hand, recently, in order to reduce the NOx in the combustion gas, a premixed lean combustion method has been developed in which the increase in flame temperature, which is the main factor of NOx generation, is suppressed and the flame temperature is lowered. According to this premixed lean combustion method, air is mixed in advance with the fuel to be supplied to the combustor liner, and the combustion temperature is made uniform, which prevents local rise in flame temperature. The amount can be reduced and the cooling performance becomes excellent. In this case, a gas turbine combustor made of a high heat resistant alloy is adopted, and this is already in practical use.

[0011]

As described above, the cooling structure of the combustor for reducing the cooling air and improving the cooling performance has produced some results up to the present. There are problems to be improved.

That is, when the heat transfer coefficient is improved for the purpose of improving the cooling efficiency, the shape of the cooling passage is simply complicated to expand the cooling area, and the cooling air becomes a non-uniform flow. This non-uniformity of flow causes non-uniform heat transfer coefficient and increase in pressure loss, and in gas turbine combustor operation, non-uniform heat transfer coefficient causes non-uniform distribution of metal temperature in the combustor. There is a problem that thermal deformation occurs in the combustor due to partial temperature rise, and partly due to the transformation point of the metal material of the combustor or the temperature rise of the inner wall surface exceeding the melting point, there is a risk of burnout. . Further, the increase in pressure loss causes an imbalance between the flow rate of cooling air that can flow in and the distribution of combustion air within the pressure difference set for each combustor, which causes problems such as impairing combustion performance.

On the other hand, the structure of the combustor liner end seal part, which is inserted and assembled inside the inlet part of the transition piece located downstream of the combustor liner for introducing the combustion gas to the gas turbine side, has an appropriate cooling air flow rate. In addition, a sealed seal type is adopted, and the mechanical joint is formed so as to hold a flexible spring force.

FIG. 10 shows a cooling structure for such a conventional combustor liner end seal portion. That is, the inlet end of the transition piece 7 is loosely fitted to the downstream end of the combustor liner 1, and a gap at this joint is formed on the outer peripheral surface of the downstream end of the combustor liner 1. It is sealed by a seal member 8 made of a curved elastic plate. In order to obtain a smooth high-speed flow of the combustion gas in the combustor, the dimensions of such a mechanical joint are such that the inner diameters of the combustor liner 1 and the transition piece 7 are substantially the same and the step difference is reduced, and the seal is sealed. It has a robust structure that facilitates insertion and withdrawal by means of the member 8, and further reduces vibration of the combustor, absorbs thermal expansion in the axial direction, absorbs mitigation of misalignment between the combustor liner 1 and the transition piece 7, and so on. It also has functions.

Therefore, it is difficult to adopt the above-described complicated cooling structure in such a portion, and even in the combustion gas temperature distribution in the axial direction of the combustor, it corresponds to the highest temperature portion, so that the cooling performance is improved. There are problems such as having to adopt a positive cooling structure.

The present invention has been made in consideration of such circumstances, and can greatly reduce the flow rate of cooling air as compared with the combustor unified by the conventional film cooling method, thus improving the cooling performance. At the same time, the premixed combustion air flow rate can be increased by increasing the temperature of the combustor and reducing the cooling air, and it is possible to obtain an ultra-lean combustion condition by suppressing the flame temperature rise, which is the main factor of NOx generation. It is an object of the present invention to provide a gas turbine combustor capable of solving both the high temperature and low NOx emission of the gas turbine combustor.

[0017]

In order to achieve the above-mentioned object, the present invention has a combustor liner forming a combustion chamber inside a combustor outer cylinder, and has an upstream end portion of the combustor liner. A pilot fuel nozzle is provided, and a multistage premixing duct for supplying a premixed gas of fuel and air is connected to the liner wall portion downstream of the pilot fuel nozzle, and the downstream end of the combustor liner is connected. A gas turbine combustor in which a transition piece for introducing combustion gas to the gas turbine side is connected, wherein the combustor liner includes an inner wall cylinder in contact with the high temperature combustion gas, and an outer wall covering the inner wall cylinder with an annular space. In a double structure consisting of a cylinder and this annular space used as a cooling air passage, a large number of air holes along the axial direction of the combustor are bored in the outer wall cylinder to impingement cool the inner wall cylinder. On the other hand, a large number of air holes, which are arranged differently from the air holes of the outer wall cylinder, are bored in the inner wall cylinder to enable film cooling of the inner wall cylinder and to support the outer peripheral side of the combustor liner. An impingement cooling portion of the inner wall cylinder is provided at a mounting portion of the support structure, and a slot type film cooling portion is provided at a seal portion with the transition piece at a downstream end of the combustor liner.

In the present invention, preferably, the impingement cooling portion at the support structure mounting portion of the inner wall cylinder of the combustor liner allows cooling air to flow through the inside of the support structure mounted to the support structure mounting portion. With a possible hollow structure, the wall thickness of the outer wall cylinder serving as the support structure mounting portion is set to be equal to or greater than that of the inner wall cylinder, and the outer wall inner cylinder is provided with a counterbore with two-step diameter air holes. Compose by

Also preferably in the present invention, the slot-type film cooling portion of the seal portion between the combustor liner and the transition piece fits between the downstream end of the outer wall cylinder of the combustion liner and the inlet end of the transition piece. A plurality of slits for elastic deformation along the axial direction of the combustor are formed in the spring seal member having an arcuate cross section inserted in the joint portion, and a cooling air inflow hole is formed in the outer wall cylindrical portion located on the inner peripheral side of the seal member. And the flow rate of the film cooling air that flows out in the same direction as the flow of the hot combustion gas, and through the slits provided in the seal member, the outer surface of the combustor liner and the inner surface of the transition piece. The flow rate of the cooling air flowing out from the formed annular gap is set to a uniform flow rate distribution that allows the flow of the cooling air within a certain pressure difference.

Further preferably in the present invention, a ring-shaped reinforcing wall is provided between the inner wall cylinder and the outer wall cylinder, and the reinforcing wall is constituted by an assembly part to the inner wall cylinder or a corrugated portion integrally molded with the inner wall cylinder. The reinforcing wall is set to have a radial thickness capable of absorbing the difference in thermal expansion between the inner wall cylinder and the outer wall cylinder and maintaining the annular space within a certain width.

[0021]

According to the gas turbine combustor of the present invention, in the main part of the combustor liner, the inner wall cylinder is strongly cooled by impingement cooling by the cooling air injected from the air holes of the outer wall cylinder to the inner wall cylinder, and further the inner wall is further cooled. The inner peripheral surface side of the inner wall cylinder is further cooled in two stages by film cooling by the cooling air blown from the air holes of the cylinder to the combustion chamber side. Therefore, since the same air flow performs the stepwise efficient cooling action, the cooling air can be reduced and the cooling performance can be improved.

Further, by providing the impingement cooling portion of the inner wall cylinder at the mounting portion of the support structure for supporting the outer peripheral side of the combustor liner, it becomes possible to efficiently cool the portion which has been difficult in the past.

In particular, the impingement cooling section at the support structure mounting portion of the inner wall cylinder of the combustor liner has a hollow structure capable of circulating cooling air inside the support structure mounted to the support structure mounting portion. At the same time, by setting the thickness of the outer wall cylinder, which is the mounting portion for the support structure, to be equal to or greater than that of the inner wall cylinder, and by piercing the outer wall inner cylinder with the air hole of the two-step diameter with the spot facing To be done.

Further, by providing the slot type film cooling portion at the seal portion with the transition piece at the downstream end of the combustor liner, efficient cooling can be performed with a relatively simple structure. The slot-type film cooling portion of the seal portion between the combustor liner and the transition piece has a bow-shaped cross-section that is inserted into the fitting portion between the downstream end of the outer wall cylinder of the combustion liner and the inlet end of the transition piece. A plurality of slits for elastic deformation along the axial direction of the combustor are formed in the spring seal member, and a cooling air inflow hole is formed in the outer wall cylindrical portion located on the inner peripheral side of the seal member, and high temperature combustion is performed. The flow rate of film cooling air discharged in the same direction as the gas flow,
The flow rate of the cooling air flowing through the slit provided in the seal member and flowing out of the annular gap formed between the outer surface of the combustor liner and the inner surface of the transition piece is set to a uniform flow rate distribution that allows inflow within a certain pressure difference. By adopting the desirable configuration, it is possible to ensure efficient cooling with a simple configuration.

Furthermore, in the present invention, a ring-shaped reinforcing wall is provided between the inner wall cylinder and the outer wall cylinder, and the reinforcing wall is constituted by an assembly part to the inner wall cylinder or a corrugated portion integrally formed with the inner wall cylinder. By configuring the reinforcing wall to have a radial thickness capable of maintaining the annular space within a certain width by absorbing the difference in thermal expansion between the inner wall cylinder and the outer wall cylinder, the configuration of each part that exerts the cooling effect described above can be obtained. It will be robust and relatively easy to implement.

[0026]

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS An embodiment of a gas turbine combustor according to the present invention will be described below with reference to the drawings. 1 is an overall sectional view schematically showing the structure of a gas turbine combustor and the flow of cooling air according to an embodiment of the present invention, and FIG. 2 is an enlarged view for explaining a portion having both impingement cooling and film cooling functions. Sectional drawing, FIG. 3 is an enlarged sectional view showing a cooling function section in a support attachment portion, and FIGS. 4 and 5 are perspective views and enlarged sectional views showing a cooling function section in a connecting section between a combustor liner and a transition piece. 6 is a sectional view showing a modified example.

As shown in FIG. 1, the gas turbine combustor of this embodiment basically has a combustor liner 13 constituting a combustion chamber 12 inside a combustor outer cylinder 11 for introducing cooling air a. A pilot fuel nozzle 14 is provided at the center of the upstream end of the combustor liner 13. Then, a multistage premixing duct 15 for supplying a lean premixed gas of fuel and air is connected to a liner wall portion on the downstream side of the pilot fuel nozzle 14, and the combustor liner 1 is further provided.
A transition piece 16 for introducing the combustion gas to the gas turbine side is connected to the downstream end of 3.

A plurality of fuels 17 for lean premixed combustion are supplied to the combustion air inlet end of each premixing duct 15, and the premixed fuel gas in each premixing duct 15 is supplied. The jet end end portion 18 is inserted into the combustor liner 4 and fixed by welding.

Then, as shown in FIG. 2, the combustor liner 1
3 has a double wall structure including an inner wall cylinder 19 that defines the combustion chamber 12 and an outer wall cylinder 20 that is arranged outside the inner wall cylinder 19 with a cylindrical space therebetween. The inner wall cylinder 19 is composed of, for example, a plurality of elements 19a arranged at intervals in the axial direction, and each element 19a is supported by the inner wall cylinder fixing portion 21 with respect to the outer wall cylinder 20 in a cantilever structure. , The thermal expansion in the axial direction of the combustor is not restricted.

It should be noted that there is a space between each of the jet outlet end portions 18 of the premixing duct 15 and the inner wall cylinder 20 of the combustor liner 13.
An annular gap 22 is formed so that the inner wall cylinder 20 and the ejection outlet cylinder end portion 18 do not come into contact with each other due to thermal expansion and thermal expansion difference.

Further, as shown in FIG. 2, a cylindrical space between the inner wall cylinder 19 and the outer wall cylinder 20 serves as an air passage 23, and the outer wall cylinder 20 is provided with a large number along the axial direction of the combustor. The air holes 24 of the inner wall cylinder 19 are impinged by the air (arrow a) guided into the combustor outer cylinder 1 from the outside of the outer wall cylinder 20 through the air holes 24 to impinge on the inner wall cylinder 19. Ment cooling.

On the other hand, the inner wall cylinder 19 also has a large number of air holes 25.
Has been drilled. The air holes 25 of the inner wall cylinder 19 are
The position of the outer wall cylinder 20 is different from that of the air hole 24, and the air (arrow b) after impingement cooling passes through the air hole 25 and is introduced into the inner wall cylinder 19, that is, the combustion chamber. ing. The air introduced from the air holes 25 of the inner wall cylinder 19 flows along the inner surface of the inner wall cylinder 19 to be used for film cooling.

A ring-shaped reinforcing wall 26 is provided on the outer peripheral surface of the inner wall cylinder 19 on the non-supporting end side so as to have a radial thickness so as not to block the air passage, and the air used for impingement cooling is supplied to the inner wall cylinder. Portion passing through the air hole 25 of 19 (arrow b)
And a portion (arrow b) passing through the outer peripheral gap 27 of the reinforcing wall 26 provided in the inner wall cylinder 19. And
The air flowing into the combustion chamber 12 is distributed in balance with each other, and these air is ejected as slot type film cooling air onto the combustor liner surface.

Incidentally, a part of the air introduced into the combustor outer cylinder 1 is introduced to the outer peripheral portions of the combustor liner 13 and the transition piece 16 and then supplied to the premixing duct 15 as combustion air.
It becomes a lean premixed fuel.

Next, the impingement cooling structure of the combustor liner support structure attachment portion 30 will be described with reference to FIG.
That is, the outer peripheral side of the combustor liner 13 has the support structure 31.
Supported by. This support structure mounting portion 30
Is a support structure 3 in which a part of the outer wall cylinder 20 is made thick.
1 has a thick outer wall cylinder 32 and a reinforcing rib 33 that reinforces the support structure 31 in the axial direction. An inner wall cylinder 19 similar to the above is provided on the inner peripheral side of the thick outer wall cylinder 32. The thick outer wall cylinder 32 has sufficient buckling resistance against external pressure and further has sufficient strength against the combustor liner supporting force in the axial direction of the combustor. The inside of the support structure 31 is hollow, and a cooling air inflow hole 34 is formed in a wall of the support structure 31.

Further, the support structure 3 is provided on the thick outer wall cylinder 32.
A plurality of air holes 35 are formed at the bottom of the cavity 1, and the air holes 35 communicate with the inside of the support structure 31 and the inflow hole 34. The air hole 35 of the thick outer wall cylinder 32 is a two-step hole with a spot facing.

The air (arrow c) flowing from the outside of the support structure 31 into the cavity through the inflow hole 34 passes through the air hole 35 of the thick outer wall cylinder 32 and flows into the air passage 23, Impingement cooling or convection cooling is performed by colliding with the inner wall cylinder 32, and then the air holes 25 of the inner wall cylinder 19
Is introduced into the combustion chamber 12 from the above and flows along the inner surface of the inner wall cylinder 19 to be used for film cooling (arrow d). Since the air hole 35 of the thick outer wall cylinder 32 is a two-step hole with a spot facing, the air passing through this portion becomes a jet flow and performs large impingement cooling or convection cooling on the inner wall cylinder 19. Since the air hole 35 has a two-step hole shape, the outer wall cylinder 2 of the combustor liner 13 is
By making the flow rate coefficient and pressure loss of the cooling air port 24 provided at 0 the same, efficient impingement cooling is enabled.

Next, cooling of the combustor liner end seal portion 40 will be described with reference to FIGS. 1, 4 and 5. 4 and 5 show the combustor liner end seal portion 40 of this embodiment.
FIG. 3 is a perspective view and an enlarged sectional view showing the slot type film cooling structure of FIG.

As shown in FIGS. 1 and 4, a spring seal member 41 having an arcuate cross section is provided at the fitting portion between the downstream end of the outer wall cylinder 20 of the combustor liner 13 and the inlet end of the transition piece 16. Has been inserted. The spring seal member 41 is connected to the outer wall cylinder 20, and the spring seal member 41, the transition piece 16 and the combustor liner 13 are mechanically joined so as to hold a flexible spring force.

Further, the spring seal member 41 is formed with a plurality of elastically deforming slits 42 along the axial direction of the combustor at intervals in the circumferential direction. A plurality of air inflow holes 43 are formed at intervals in the circumferential direction in the outer wall cylinder 20 located on the inner peripheral side of the spring seal member 41, and a bias plate made of a thin plate is formed on the inner peripheral surface side of the air inflow holes 43. 44 are provided. An annular gap 45 communicating with the combustion chamber 12 side through the outer surface of the combustor liner and the inner surface of the transition piece is formed behind the slit 42 of the spring seal member 41.

The air inside the combustor outer cylinder 11 is slit 42.
The film cooling air (arrow e) which flows into the combustion chamber 12 through the air inflow hole 43 in the inner peripheral side of the spring seal member 41 through the ), And flows into the inner peripheral side of the spring seal member 41 through the slit 42, and then comes out of the slit 42 again.
1 is divided into air (arrow f) flowing into the combustion chamber 12 from the annular gap 45 via the outer peripheral side of 1. The flow rate of these cooling air is set to a uniform flow rate distribution that allows the flow into the combustion chamber 12 within a constant pressure difference.

The cooling air is used as film cooling air in the above-mentioned path, and cools the inner surface of the combustor on the downstream side of blowing. As a result, it is possible to efficiently lower the combustor liner metal temperature by using the slot-type film cooling and the external surface cooling together in the combustor liner end seal portion which has conventionally been only the external surface cooling.

In the present embodiment described above, since the metal temperature distribution of the inner wall cylinder 19 of the combustor liner 13 due to the cooling air can be set so as to have a uniform and optimum heat transfer coefficient distribution, the cooling air flow rate is reduced, The amount of combustion air required for premixed combustion can be increased, and the premixed fuel-air ratio can be set to a combustion temperature at which NOx is not produced.

As described above, according to the gas turbine combustor according to the present embodiment, the cantilever double cylinder type cooling structure for both impingement cooling and film cooling arranged in multiple stages in the axial direction of the combustor is adopted. Since the combustor liner end seal portion 40 adopts the slot type film cooling and the combustor liner support structure mounting portion 30 adopts the impingement cooling, it is compared with the combustor unified by the conventional film cooling method. It is possible to easily reduce the cooling air flow rate to about 60% of the conventional level, improve the cooling performance, and at the same time increase the premixed combustion air flow rate by raising the temperature of the combustor and reducing the cooling air. , It is possible to obtain the ultra-lean combustion condition by suppressing the flame temperature rise that is the main factor of NOx generation, which contributes to the reduction of NOx and at the same time makes the metal temperature distribution uniform. It has to. That is, there is an excellent effect that both high temperature and low NOx of the gas turbine combustor can be solved.

Further, according to the gas turbine combustor of this embodiment, in the main part of the combustor liner 13, the inner wall cylinder 19 is impingement-cooled by the cooling air injected from the air holes 24 of the outer wall cylinder 20 to the inner wall cylinder 19. Is strongly cooled, and the inner peripheral surface side of the inner wall cylinder 19 is further cooled in two stages by film cooling by the cooling air blown from the air holes 25 of the inner wall cylinder 19 to the combustion chamber 12 side. Therefore, since the same air flow performs the stepwise efficient cooling action, the cooling air can be reduced and the cooling performance can be improved.

Further, by providing the impingement cooling portion having the inner wall cylinder in the support structure mounting portion 30 for supporting the outer peripheral side of the combustor liner, it becomes possible to efficiently cool the portion which has been difficult in the past. In particular, the impingement cooling portion of the support structure attachment portion 30 of the inner wall cylinder 19 of the combustor liner 13 and the inside of the support structure 31 attached to the support structure attachment portion 30 have a hollow structure capable of circulating cooling air. In addition, by setting the wall thickness of the outer wall cylinder 20 serving as the support structure attachment portion to be equal to or greater than that of the inner wall cylinder 19, and by boring the outer wall inner cylinder 19 with the counterbore two-step diameter air hole 43. The above action is more reliably performed.

Further, in the combustor liner end seal portion 40, which is a seal portion with the transition piece 16 at the downstream end portion of the combustor liner 13, by providing a slot type film cooling portion, a relatively simple structure is provided. Can perform efficient cooling.
The slot-type film cooling portion of the combustor liner end seal portion 40 has an arcuate cross-sectional shape inserted into a fitting portion between the downstream end portion of the outer wall cylinder 20 of the combustor liner 13 and the inlet end portion of the transition piece 16. A plurality of slits 42 for elastic deformation along the axial direction of the combustor are formed in the spring seal member 41, and a cooling air inflow hole 43 is formed in the outer wall cylindrical portion 20 located on the inner peripheral side of the seal member. In addition, the flow rate of the film cooling air flowing out in the same direction as the flow of the high temperature combustion gas, and the outer surface of the combustor liner 13 and the inner surface of the transition piece 16 are formed by passing through the slit 42 provided in the spring seal member 41. By setting the flow rate of the cooling air flowing out from the annular gap 45 that is set to a uniform flow rate distribution that allows inflow within a constant pressure difference,
With a simple structure, efficient cooling can be surely performed.
For example, compared with a combustor unified by a conventional film cooling method, it has become possible to reduce the cooling air flow rate to about 60% of the conventional one.

FIG. 6 shows a modification of the inner wall cylinder 19. In this example, the reinforcing wall 46 is formed by a corrugated portion formed by integrally molding the inner wall cylinder 19. The difference in thermal expansion between the inner wall cylinder 19 and the outer wall cylinder 20 is absorbed by the reinforcing wall 46, and the air passage 23 formed by the annular space is set to have a radial thickness that can be maintained within a certain width.

By thus forming the reinforcing wall 46 as a corrugated cross-section portion integrally formed with the inner wall cylinder 19,
The inner wall cylinder 19 and the outer wall cylinder 20 in consideration of thermal expansion of the inner wall cylinder 19 form a gap at a constant interval. Therefore, the processing cost of the inner wall cylinder 19 can be reduced, and the configuration of each part having a cooling effect can be realized firmly and relatively easily.

[0050]

As described in detail in the above embodiments, according to the gas turbine combustor according to the present invention, a cantilever double cylinder for both impingement cooling and film cooling arranged in multiple stages in the axial direction of the combustor. The same type of cooling structure was adopted, the slot type film cooling was adopted for the combustor liner end seal part, and the impingement cooling was adopted for the combustor liner support structure mounting part. It is possible to easily reduce the cooling air flow rate significantly compared to a combustor, and improve the cooling performance, while at the same time increasing the temperature of the combustor and increasing the premixed combustion air flow rate by reducing the cooling air. Therefore, it is possible to suppress the flame temperature rise, which is the main factor of NOx generation, and obtain an ultra-lean combustion condition, which raises the temperature of the gas turbine combustor and lowers NOx.
The excellent effect of being able to solve both sides of the problem is achieved.

[Brief description of drawings]

FIG. 1 is an overall sectional view schematically showing the structure of a gas turbine combustor and the flow of cooling air according to an embodiment of the present invention.

FIG. 2 is an enlarged cross-sectional view for explaining a portion having both impingement cooling and film cooling functions in an embodiment of the present invention.

FIG. 3 is an enlarged cross-sectional view showing a cooling function part of a support attachment part in one embodiment of the present invention.

FIG. 4 is a perspective view showing a cooling function portion in a connecting portion between the combustor liner and the transition piece according to the embodiment of the present invention.

FIG. 5 is an enlarged cross-sectional view showing a cooling function part in a connecting part between the combustor liner and the transition piece in the embodiment of the present invention.

FIG. 6 is a sectional view showing a modified example of the embodiment of the present invention.

FIG. 7 is a sectional view showing a conventional example.

FIG. 8 is a cross-sectional view showing another conventional example.

FIG. 9 is a sectional view showing still another conventional example.

FIG. 10 is a sectional view showing still another conventional example.

[Explanation of symbols]

 11 Combustor Outer Cylinder 12 Combustion Chamber 13 Combustor Liner 14 Pilot Fuel Nozzle 15 Premixing Duct 16 Tail Cylinder 17 Fuel 18 Outlet Cylinder End 19 Inner Wall Cylinder 20 Outer Wall Cylinder 21 Cylinder Fixed Part 24 Air Hole 25 Air Hole 26 Ring-shaped Reinforcement wall 27 Peripheral clearance 30 Liner support structure mounting portion 31 Support structure 32 Thick outer wall cylinder 33 Reinforcing rib 34 Air inflow hole 35 Air hole 41 Spring seal member 42 Slit 43 Air inflow hole 44 Drift plate

Claims (4)

[Claims] 1. A combustor liner forming a combustion chamber is arranged inside a combustor outer cylinder, a pilot fuel nozzle is provided at an upstream end of the combustor liner, and a pilot fuel nozzle is provided downstream of the pilot fuel nozzle. Gas connected to a multi-stage premixing duct for supplying a premixed gas of fuel and air to the liner wall portion, and a tail pipe for introducing combustion gas to the gas turbine side at the downstream end of the combustor liner A turbine combustor, wherein the combustor liner has a double structure including an inner wall cylinder in contact with a high temperature combustion gas and an outer wall cylinder covering the inner wall cylinder with an annular space open, and the annular space is provided with a cooling air passage. In the above, the outer wall cylinder is provided with a large number of air holes along the axial direction of the combustor to allow impingement cooling of the inner wall cylinder, while the inner wall cylinder has air holes of the outer wall cylinder. And the inner wall cylinder can be film-cooled by forming a large number of air holes differently arranged, and the impingement cooling portion of the inner wall cylinder is attached to the mounting portion of the support structure that supports the outer peripheral side of the combustor liner. A gas turbine combustor, further comprising a slot-type film cooling section provided at a seal portion with the transition piece at a downstream end of the combustor liner. 2. The gas turbine combustor according to claim 1, wherein the impingement cooling portion at the support structure attachment portion of the inner wall cylinder of the combustor liner is provided inside the support structure attached to the support structure attachment portion. In addition to having a hollow structure through which cooling air can flow, the thickness of the outer wall cylinder serving as the support structure mounting portion is set to be equal to or greater than that of the inner wall cylinder, and the outer wall inner cylinder has a counterbore with a two-step diameter. A gas turbine combustor characterized by being constructed by forming holes. 3. The gas turbine combustor according to claim 1, wherein the slot-type film cooling portion of the seal portion between the combustor liner and the transition piece is the downstream end of the outer wall cylinder of the combustion liner and the inlet of the transition piece. A plurality of slits for elastic deformation along the axial direction of the combustor are formed in a spring seal member having an arcuate cross-section inserted in a fitting portion with an end portion, and an outer wall cylindrical portion located on the inner peripheral side of the seal member. And a flow rate of the film cooling air that flows out in the same direction as the flow of the high temperature combustion gas, and the outer surface of the combustor liner passing through the slit provided in the seal member. A gas turbine combustor characterized in that the flow rate of cooling air flowing out from an annular gap formed on the inner surface of the transition piece is set to be a uniform flow rate distribution that allows inflow within a constant pressure difference. 4. The gas turbine combustor according to claim 1, wherein a ring-shaped reinforcing wall is provided between the inner wall cylinder and the outer wall cylinder, and the reinforcing wall is integrally formed with an assembly part of the inner wall cylinder or the inner wall cylinder. A gas turbine comprising a corrugated portion, wherein the reinforcing wall is set to have a radial thickness capable of absorbing a difference in thermal expansion between the inner wall cylinder and the outer wall cylinder to maintain the annular space within a certain width. Combustor.