JP4123120B2 - Gas turbine combustor - Google Patents

Description

  The present invention relates to a gas turbine combustor, and more particularly, it reduces the flow rate of air that passes through a mating connection portion between a combustion chamber liner and a transition piece, and at the same time suppresses a rise in metal temperature of the combustion chamber liner, thereby reducing nitrogen oxide (hereinafter referred to as nitrogen oxide) The present invention relates to a gas turbine combustor capable of reducing the emission amount of NOx) and ensuring the reliability of the combustor.

  A reduction in the amount of NOx emitted when the fuel is burned is important for suppressing an increase in environmental load. Since NOx emissions greatly depend on the combustion flame temperature, it is effective to decrease the flame temperature by increasing the flow rate of combustion air in the main combustion reaction field to achieve lean combustion. In addition, when premixed combustion is performed in which air and fuel are mixed in a premixer and then supplied to the combustion chamber for combustion, NOx can be greatly reduced as compared with diffusion combustion in which fuel and air are mixed and burned. In the case of premixed combustion, an increase in the combustion air flow rate is more effective in reducing NOx emissions.

  Under such circumstances, in gas turbine engines, proposals have been made to use as much combustion air as possible from the compressor into the combustor. This leads to a reduction in the cooling air flow rate of the members constituting the combustor that does not substantially contribute to the combustion reaction. As a method for cooling the combustion chamber liner, a method of forming film cooling air on the inner wall of the combustion chamber liner by providing a number of slots on the outer wall of the combustion chamber liner is generally employed. There are many suggestions for

  Similarly, it is possible to increase the flow rate of combustion air by reducing the flow rate of air that slips through a mating connection portion with a transition piece or a duct called a communication duct disposed between the combustion chamber liner and the turbine. Become. Here, as the fitting connection structure, one end of an elastic plate formed on an arc having a slit called a spring seal or a hula seal is spot welded at the downstream end of the main chamber liner, and the spring seal portion is transitioned. The method of pushing inside the piece is adopted. By doing so, the spring seal that is in close contact with the inside of the transition piece during the assembly is elastically deformed, and the mating portions of the main chamber liner and the transition piece are not fixed, so that they can slide relative to each other. Can absorb thermal elongation. In order to reduce the flow rate of air passing through this part, the gap between the spring seal and the transition piece in the assembled state is reduced, or the width and number of slits of the spring seal are reduced to reduce the opening area. To achieve.

  The combustion chamber liner structure by film cooling related to the present invention is described in, for example, Patent Document 1 and Patent Document 2. Patent Document 3 discloses an example of a spring seal structure, Patent Document 4 discloses an example of a means for reducing the air flow rate that slips through the fitting connection portion between the combustion chamber liner and the transition piece, and Patent Document 5 describes a combustion chamber liner. An example of means for cooling the mating connection structure with the transition piece is described.

Japanese Patent No. 2596921 JP 2001-289062 A JP 2002-89843 A JP-A-10-153129 Japanese Patent Laid-Open No. 10-38276

  In the present invention, there is a flow path through which air flows outside the combustion chamber liner and the transition piece, and the flow direction of the air flowing through the flow path is from the transition piece to the combustion chamber liner, and is generated inside the combustion chamber liner. The present invention provides a means for improving the reliability of the combustion chamber liner suitable for a so-called return flow type gas turbine combustor configured such that the combustion gas is guided to the turbine through a transition piece. In the return flow type gas turbine combustor, a transition piece is cooled by disposing a sleeve outside the transition piece and using an air flow path as a gap with the sleeve. On the other hand, also in the combustion chamber liner, a sleeve or an outer cylinder is arranged on the outer side, and a gap between the sleeve or the outer cylinder is used as an air flow path. The air flowing through this air flow path also contributes to the cooling of the combustion chamber liner. However, in the case of the combustion chamber liner, high-temperature combustion gas is locally formed depending on the combustion mode, so that further cooling enhancement may be required. It is common. As this means, a method of forming film cooling air on the inner wall of the combustion chamber liner by providing a large number of slots on the outer wall of the combustion chamber liner is generally employed.

  As described above, a spring seal is generally used as a fitting connection structure between the combustion chamber liner and the transition piece. In this case, the radial flow path width of the air flow path at the fitting connection position is on the transition piece side. Smaller and larger on the combustion chamber liner side, and a step is generated in the radial direction at that portion. In addition, as described above, the reduction of the air flow rate that slips through the mating connection portion between the combustion chamber liner and the transition piece can be achieved by reducing the gap between the spring seal and the transition piece in the assembled state, or by the slit that the spring seal has. This is achieved by means such as narrowing the width of.

  The present invention grasps a problem that appears when the flow rate of air passing through the fitting connection portion between the combustion chamber liner and the transition piece is reduced by the above-described means, and provides a simple and effective means for this problem. It was made in the process of finding out.

  That is, the air flowing outside the transition piece has a high flow velocity in order to ensure a cooling effect, and in this state, reaches the fitting connection portion between the combustion chamber liner and the transition piece. In the mating connection portion, there is a step in the radial direction in the air flow path, so that a circulation flow region is formed. The flow rate of air passing through the mating connection between the combustion chamber liner and the transition piece is affected by this circulating flow, but when the metal temperature at the downstream end of the combustion chamber liner is measured, there is a large deviation in the circumferential distribution of the metal temperature. It was found that the state of the circulation flow is not uniform in the circumferential direction. This is due to the fact that the flow velocity distribution of the air ejected from the transition piece to the combustion chamber liner is not uniform in the circumferential direction due to the influence of the shape of the transition piece and the like. The combustion chamber liner must be free from the appearance of local high metal temperature sites in order to ensure reliability. However, in order to prevent the appearance of a high metal temperature region locally, the gap between the spring seal and the transition piece is increased, or the opening area is increased by increasing the width and number of slits of the spring seal. , Increasing the air flow rate that slips through the mating connection between the combustion chamber liner and the transition piece deviates from the viewpoint of increasing the combustion air flow rate as much as possible to reduce NOx emissions. .

  The present invention can reduce the air flow rate that has passed through the fitting connection between the combustion chamber liner and the transition piece by a relatively simple method, and can make the circumferential distribution of the cooling air flow rate of the combustion chamber liner substantially uniform. An object is to provide a combustor.

  In order to solve the above-mentioned problems, the present invention provides a gas turbine combustor structure in which a combustion chamber liner and a transition piece are connected with a fitting structure, and an air flow path is provided outside the both. A partition wall that provides resistance to air flow in the vicinity of the end of the transition piece at the fitting connection position so that a step in which the radial flow path width of the air flow path is small on the transition piece side and large on the combustion chamber liner side is generated. A ring having a height ranging from 0.8 times the level difference of the flow path to about 60% of the radial flow path width in the radial direction of the air flow path on the combustion chamber liner side is provided on the outer wall of the combustion chamber liner.

  Further, according to the present invention, in the gas turbine combustor described above, a dilution hole is provided in the combustion chamber liner, and air flow resistance is provided between the end of the transition piece on the fitting connection position side and the dilution hole arranged in the combustion chamber liner. A ring is provided on the outer wall of the combustion chamber liner.

  According to the present invention, it is possible to reduce the air flow rate that has passed through the fitting connection between the combustion chamber liner and the transition piece, and to make the circumferential distribution of the cooling air flow rate of the combustion chamber liner uniform by a relatively simple method. Therefore, an increase in the combustion air flow rate that leads to a reduction in the NOx emission amount can be achieved, and at the same time, an increase in the metal temperature at the end of the combustion chamber can be suppressed.

  An embodiment of a gas turbine combustor according to the present invention will be described with reference to FIG. A combustion chamber liner according to the present invention, in which a ring serving as an air flow resistance is provided on the outer wall near the end of a transition piece, is used by being incorporated in a gas turbine combustor as shown in FIG.

  A plurality of cans of the gas turbine combustor of the present embodiment are arranged around a gas turbine (not shown). Each combustor can has an outer cylinder 1, a combustion chamber liner 2, and a transition piece 3. A sleeve 4 is disposed outside the transition piece 3. A part of the sleeve 5 is disposed outside the combustion chamber liner 2.

  Combustion air 50 is supplied from a compressor (not shown), passes between the transition piece 3 and the sleeve 4, passes between the combustion chamber liner 2, the sleeve 5, and the outer cylinder 1 and reaches the diffusion burner 10.

  The combustion air 50 and the fuel 60 are combusted inside the combustion chamber liner 2, and the generated combustion gas 70 is sent to a turbine (not shown) through the transition piece 3 to drive the turbine.

  In FIG. 2, the enlarged view of the fitting connection part of the combustion chamber liner 2 and the transition piece 3 is shown. The combustion chamber liner 2 and the transition piece 3 have a fitting structure with a spring seal 101 interposed therebetween. As shown in FIG. 3, the spring seal 101 has a structure in which an elastic plate is formed in an arc shape and one end thereof is spot welded to the combustion chamber liner 2. In the embodiment of FIG. 3, the spring seal 101 is provided with a slit 102 having a width of about 1 mm. One end of the slit 102 reaches the end of the spring seal. The through air 51 shown in FIGS. 2 and 4 flows through the slit 102 to the end 400 of the combustion chamber liner.

  In the fitting connection portion between the conventional combustion chamber liner 2 and the transition piece 3, since there is a radial step 300 in the air flow path, a circulation flow region 200 B as shown in FIG. 6 is formed by the jet of combustion air 50. It is formed. Since this circulation flow region 200B is formed in contact with the step 300, it is not a sufficient condition for recovering the dynamic pressure of the jet of the combustion air 50 and increasing the static pressure. The step 300 and the end 400 of the combustion chamber liner The total pressure difference between was small. However, in the conventional spring seal 101 of the combustion chamber liner 2, the opening area of the slit 102 is set large so that a sufficient flow rate of the slip-through air 51 is secured even at this total pressure difference. Thus, although there was a variation in the metal temperature in the circumferential direction, the metal temperature at the end of the combustion chamber liner was well within the allowable range.

  However, when the opening area of the slit 102 of the spring seal 101 is reduced in order to reduce the NOx emission amount, as shown in FIG. 5, there is a portion where the metal temperature rises as a whole and becomes locally high. Appeared.

In this embodiment, in order to solve this problem, as shown in FIG. 2, a ring 100 serving as an air flow resistance is disposed on the outer wall of the combustion chamber liner 2 in the vicinity of the transition piece end 301 at the fitting connection position. . Here, it is desirable that the ring 100 be disposed away from the transition piece end 301 more than the distance corresponding to the outside of the circulation flow region 200B when the ring 100 shown in FIG. 6 is not disposed. The range of the circulation flow region 200B can be grasped relatively easily by the recent development of flow analysis technology. The height of the ring 100 is preferably in a range from 0.8 times the step 300 of the air channel to about 60% of the radial channel width on the combustion chamber liner side. When the ring 100 is thus arranged on the outer wall of the combustion chamber liner 2, the core region 200 </ b> A of the circulating vortex is formed away from the transition piece end 301. As a result, the dynamic pressure of the jet of the combustion air 50 is restored and the static pressure is increased, and the total pressure difference between the step 300 and the end 400 of the combustion chamber liner is increased, so the flow rate of air passing through the spring seal 101 is increased. To do. If the height of the ring 100 is smaller than 0.8 times the step 300 of the air flow path, the recovery of the dynamic pressure of the jet of the combustion air 50 is not sufficient, and the radial direction of the air flow path on the combustion chamber liner side If the height is equal to or greater than about 60% of the flow path width, the pressure loss of the combustor increases due to the flow path resistance, which adversely affects the thermal efficiency of the gas turbine engine. By setting the height of the ring 100 in the above range, the metal temperature at the end of the combustion chamber was reduced substantially uniformly as shown in FIG. 5, and the variation width of the metal temperature was also reduced. In the present embodiment, the variation width of the metal temperature is reduced as compared with the conventional example. The ring 100 makes the formation state of the circulating vortex more uniform in the circumferential direction than the conventional example, and the air flow rate passing through the spring seal 101 is reduced. This is because it is more uniform in the circumferential direction than the conventional example. Therefore, this embodiment can achieve the intended purpose of increasing the combustion air flow rate as compared with the conventional example. Considering the operation of this embodiment, the shape of the ring 100 is not particularly limited. Although the shape shown in FIG. 2 is L-shaped, a flat ring may be used, and the thickness is not limited. It is also permissible to change the shape in the circumferential direction so that the combustion chamber liner can be easily incorporated, as long as the above action is not significantly impaired.

  According to the present embodiment, even if the flow rate of air passing through the fitting connection portion between the combustion chamber liner and the transition piece is reduced, the rise in the metal temperature at the end of the combustion chamber liner can be suppressed, so the reliability of the combustion chamber liner There is an effect that can be secured.

  FIG. 4 shows another embodiment of the present invention. The difference from the embodiment of FIG. 2 is that a dilution hole 401 is provided in the combustion chamber liner 2. The action and effect of the dilution hole 401 are well-known matters for engineers working on gas turbine combustors. In the present embodiment, the ring 100 serving as an air flow resistance is disposed between the transition piece end 301 and the dilution hole 401. If the ring 100 is disposed at a position where the distance from the transition piece end 301 is larger than the dilution hole 401, the circulating vortex formed between the step 300 and the ring 100 may enter the combustion chamber liner 2 from the dilution hole 401. The uniformity in the circumferential direction is significantly impaired due to the flow rate of the ejected air. As a result, the metal temperature at the end of the combustion chamber liner 2 cannot be cooled with a small air flow rate. The cooling of the end of the combustion chamber liner 2 can be effectively realized by disposing the ring 100 between the transition piece end 301 and the dilution hole 401 as in this embodiment.

  According to the present embodiment, even if the dilution hole is arranged in the combustion chamber liner, there is an effect that the end of the combustion chamber can be cooled with a small air flow rate.

1 is a cross-sectional view of a gas turbine combustor according to an embodiment of the present invention. Sectional drawing which expanded a part of gas turbine combustor regarding one Example of this invention. Sectional drawing which expanded a part of gas turbine combustor. Sectional drawing which expanded a part of gas turbine combustor regarding the other Example of this invention. The figure explaining the comparison of the metal temperature distribution of the combustion chamber end. Sectional drawing which expanded a part of conventional gas turbine combustor.

Explanation of symbols

DESCRIPTION OF SYMBOLS 1 ... Outer cylinder, 2 ... Combustion chamber liner, 3 ... Transition piece, 4, 5 ... Sleeve, 10 ... Diffusion burner, 20 ... Premix burner, 50 ... Combustion air, 60 ... Fuel, 70 ... Combustion gas,
DESCRIPTION OF SYMBOLS 100 ... Ring, 101 ... Spring seal, 300 ... Level difference, 400 ... End of combustion chamber liner, 401 ... Dilution hole.

Claims (4)

A gas turbine combustor configured to combust air and fuel pressurized by a compressor in a combustor liner, and to guide the generated combustion gas to the turbine through the transition piece. The transition piece and the combustor liner Resistance of air flow flowing to the combustor side downstream of the combination part of the transition piece and the combustor liner in a flow path for guiding air from the outer wall surface to the combustor surrounded by a guide wall installed at a predetermined interval And a step is formed in which the radial flow path width of the flow path for guiding the air is small on the transition piece side and large on the combustion chamber liner side, and the partition wall is zero of the step of the air flow path. The height is in a range from 8 times to 60% of the radial flow path width in the radial direction of the air flow path on the combustion chamber liner side. A gas turbine combustor which is provided on the outer wall surface side of the combustor liner in the vicinity of the end of the transition piece at the fitting connection position of the combustor piece and the combustor liner .   2. The static pressure downstream of the combination part of the transition piece and the combustor liner is increased by providing resistance to the air flow flowing to the combustor by the partition wall, and flows into the combination gap between the transition piece and the combustor liner. A gas turbine combustor having an increased air flow rate.   The gas turbine combustor according to claim 1, wherein the partition wall is provided on an outer wall surface side of the combustor liner. 4. The gas according to claim 3, wherein the combustion chamber liner is provided with a dilution hole for mixing combustion gas with air, and the partition is provided between the end of the transition piece and the dilution hole at the combination location. Turbine combustor.

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