JPS63243631A - Gas turbine combustor cooling structure - Google Patents

Abstract

PURPOSE:To reduce or eliminate problems such as heat exposure damage, cracks, shorter life, etc. of a combustor by forming axially parallel inner channels on the inner wall of the outer cylinder, and disposing communicating holes which have a diameter larger than the inner channel width, near the outer cylinder upstream end and the inner cylinder downstream end coaxially with the inner channels. CONSTITUTION:Each of multiple cylinders 3-6 having an increasingly larger diameter has a double structure in which an outer cylinder 12 and an inner cylinder 13 are overlapped. Multiple inner channels 14 having a rectangular section are formed on the inner wall of the outer cylinder in parallel with the combustor axis. Further, multiple cooling air communicating holes 15, 16 having a diameter larger than the width of the inner channels 14 are drilled in a staggered manner circumferentially near the upstream end of the outer cylinder 12 and the downstream end of the inner cylinder 13 coaxially with the inner channels 14. Because of this, a uniform cooling can be performed over the entire axial length of the combustor inner cylinder 13 by means of air cooling or film cooling. Therefore, problems of heat exposure damage, crack, shorter life, etc., of the combustor due to the increased output and temperature can be reduced or eliminated.

Description

[Detailed description of the invention] Industrial applications The present invention relates to the structure of a combustor inner cylinder of a gas turbine.

2. Description of the Related Art Conventional examples of cooling structures for combustors in industrial and aircraft gas turbines include those shown in FIGS. 7 and 8.

This combustor (inner cylinder) l is designed to take in cooling air into the gap between a plurality of cylinders 2 to 6, whose inner diameters gradually increase from upstream to downstream, overlapped at their adjacent ends. Partition walls, i.e. wavy louvers
r) 7 is inserted.

FIG. 7 shows a case where the number of cylinders is five, for example. Taking cylinders 2 and 3 as an example, the louver 6
The cylinders 2 and 3 are made into an integral structure by being joined at portions A and B that contact the inner wall surface of the cylinder 3 and the outer wall surface of the cylinder 2, respectively. And the other cylinder 4~
6, the combustor 1 consisting of one inner cylinder is formed by continuously joining louvers 6, which are the same means, into the gaps between the respective ends.

As shown in FIG.
The inner wall surface on the downstream side of each cylinder 2 to 6 is cooled by flowing toward the downstream side inside the combustor 1 through the cooling air passages C and D defined by the cylinders 2 to 6. .

This cooling method is a well-known technique and is generally called a film cooling method.

On the other hand, the fuel provided for combustion in the combustor 1 is
Combustion is performed by mixing with primary air that flows into the combustor (inner cylinder) from an injection nozzle 8 provided on the upstream side of the combustor l, that is, in the cylinder 1.

Furthermore, the air is connected to the downstream side of the combustor 1 by remixing it with secondary air flowing in from the secondary air supply port 11 which is also opened on the downstream side of the combustor 1, that is, mainly on the circumference of the cylindrical part 6. The gas temperature is reduced to maintain a suitable combustion gas temperature at the inlet of a turbine (not shown).

Problems to be Solved by the Invention The conventional gas turbine combustor cooling structure described above has the following problems:
However, there were the following problems.

With the recent increase in the output of gas turbines, the temperature of the gas at the exit of the combustor has increased, so the metal temperature of the inner cylinder forming the combustor has tended to increase.

However, even if a rare and expensive super heat-resistant alloy plate is adopted to cope with this high temperature, when applying the conventional cooling method (film cooling), the entire length of the combustor inner cylinder is
There was a strong tendency that the downstream side of each cylinder 2 to 6 became high temperature without being cooled on average, while the temperature near the upstream side where these louvers 7 were provided was too low.

Therefore, on the wall surfaces of these cylinders 2 to 7, it is no longer possible to avoid burnout, occurrence of cracks, and shortened lifespan.

Means for Solving the Problems In order to solve these conventional problems, the present invention provides the following:
In a gas turbine combustor, which forms a single inner cylinder by connecting the adjacent ends of a plurality of cylinders whose inner diameters gradually increase, these cylindrical parts have a double structure in which an outer cylinder and an inner cylinder are overlapped. A large number of inner grooves extending parallel to the axial direction of the combustor inner cylinder are provided on the inner wall surface of the outer cylinder, and along the same axis of each of these inner grooves, a groove is formed upstream of each of the outer cylinders. A large number of communication ports having a diameter larger than the width of the inner groove are bored in a staggered manner in the circumferential direction near the end portions of the side and the downstream side of each inner cylinder.

According to this method, each cylindrical part forming the inner cylinder of the combustor has a double structure in which the outer cylinder and the inner cylinder are overlapped, and the outer cylinder is provided with an inner groove, and the inner grooves of the brackets have the same inner groove. Since communication ports are provided along the axis in the circumferential direction near the upstream ends of the outer tube and the downstream ends of the inner tube, cooling air can flow into the inner grooves.

EXAMPLE Hereinafter, an example of a combustor cooling structure for a gas turbine according to the present invention will be described in detail with reference to FIGS. 1 to 6. In these figures, the same parts as shown in FIGS. 7 and 8 are designated by the same reference numerals, and detailed explanation thereof will be omitted.

According to the present invention, as shown in FIGS. 1 to 3,
Each of the plurality of cylinders 3 to 6 whose inner diameter increases successively is formed by mechanical means, etc. ( (not shown).

Furthermore, as shown in FIGS. 4 and 5, along the same axis of each of these inner grooves, the upstream side of each outer cylinder 12 (see FIG. 4)
A large number of cooling air communication ports 15 and 16 having a diameter sufficiently larger than the width of the inner groove 14 are arranged in a staggered manner in the circumferential direction near the respective ends of the downstream side of each inner cylinder 13 (see FIG. 5). to be drilled. The inner diameter of these communication ports is preferably at least twice the width of the inner groove 14.

As shown in FIG. 6, in the adjacent ends of each of the cylindrical parts 3 to 6 having such a structure, the outer cylinder 12 of the cylinder with the smaller inner diameter (for example, cylinder 3) and the inner diameter The inner cylinder 13 of the larger cylinder (for example cylinder 4) may be circumferentially welded. Moreover, even if these outer cylinder 12 and inner cylinder 13 are each extended parallel to the axial direction of the combustor (inner cylinder) 1 and joined at the extension part (not shown), good. As a result, a combustor l consisting of one inner cylinder is formed. In addition, in the cylinder 2, the heating temperature of the wall surface due to combustion is usually higher than that of the other cylinders 3 to 6.
The communication ports 15 and 16 do not necessarily need to be provided, since they are not expensive compared to the above.

With the above configuration, the cooling air introduced from the outside of the combustor 1 in stages flows in from the communication port 15 of the outer cylinder 12 of each cylinder 3 to 6 shown in FIG. After passing 1M14 of the large number from upstream to downstream, the third
It is discharged into the combustor through the communication port 16 of the inner cylinder 13 shown in the figure.

In the process of such a cooling air flow, by flowing the cooling air into the inner groove 14, not only each inner cylinder 13 exposed to high temperature combustion gas in the combustor 1 is sufficiently cooled, but also the inner cylinder 13 is cooled at the same time. By discharging cooling air into the combustor l from the communication port 16 of the combustor l, the effect of film cooling can also be performed synergistically.

As described above, air cooling and film cooling are applied to each cylinder 3 to 6.
By performing the cooling in each section, uniform cooling can be efficiently performed overall over the entire axial length of the combustor (inner cylinder) (2).

Effects of the Invention As detailed above, according to the present invention, uniform cooling can be performed by air cooling and film cooling over the entire axial length of the inner cylinder during combustion 2, so that high temperature due to increased output of the gas turbine can be achieved. It is possible to change.

Moreover, as a result of this combustor improvement, the reliability and performance of the high temperature gas turbine system can be increased.

[Brief explanation of drawings]

FIG. 1 is a longitudinal cross-sectional view of a main part showing an example of the combustor cooling structure of a gas turbine according to the present invention, FIG. 2 is a perspective view taken along the line ■-■ mainly showing the outer cylinder side of FIG. 1, and FIG. Fig. 1 is a perspective view taken along the line ■-■ mainly showing the inner cylinder side; Fig. 4 is a sectional view along the vertical line ■-■ in Fig. 2; and Fig. 5 is a cross-section taken along the vertical line ■-V in Fig. 3. Fig. 6 is an assembled overview diagram showing the entire combustor inner cylinder, Fig. 7 is an assembled overview diagram showing the combustor cooling structure of a conventional gas turbine, and Fig. 8 is a part of the joint of the louver in Fig. 7. FIG. l...Combustor (inner cylinder), 2-6...Cylinder, 12...Outer cylinder, 13...Inner cylinder, 14...Inner groove, 15.16...Figure 4

Claims (1)

[Claims] In a gas turbine combustor that forms one inner cylinder by connecting the adjacent ends of a plurality of cylinders whose inner diameters gradually increase, each of these cylinders is constructed as a double structure in which an outer cylinder and an inner cylinder are overlapped. , a large number of inner grooves extending parallel to the axial direction of the combustor inner cylinder are provided on the inner wall surface of the outer cylinder, and further, along the same axis of each of these inner grooves, a groove is formed on the upstream side of each outer cylinder. and a combustor cooling structure for a gas turbine, in which a large number of communication ports having a diameter larger than the width of the inner groove are bored in a staggered manner in the circumferential direction near the downstream end of each inner cylinder.