JPH11257003A - Impingement cooling device - Google Patents

Abstract

PROBLEM TO BE SOLVED: To prevent generation of a cross flow and to improve cooling performance by defining a cavity in a high temperature member of a gas turbine with an impingement plate, and dividing the cavity into a plurality of chambers with the impingement plate, and forming a clearance with throttled path between each cavity. SOLUTION: A partition 17 is provided in width direction at almost center of a stationary shroud 13 of a gas turbine. A first impingement plate 11 made of a heat resisting metal and having many cooling vents drilled therein is fixed on frame portions 13a and 13b of the shroud 13 by welding both ends thereof. A second impingement plate 20 is disposed by fixing one end thereof to an upper end edge of the partition 17 and the other end to the frame 13b. Accordingly, at the upper portion of the shroud 13, a first cavity 16 is formed between a bottom plate 13c of the shroud 13 and the first impingement plate 11, a second cavity 19 is formed between the first and the second impingement plates 11 and 20. The cavities 16 and 19 are communicated via a clearance 18 with throttled path.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to an impingement cooling device for cooling a high-temperature member, that is, impingement cooling, through a plurality of impingement cooling holes (small holes) formed in an impingement plate by using a cooling medium.

[0002]

2. Description of the Related Art High temperature turbine components such as blades and stationary blade shrouds of gas turbines are used under extremely high gas temperature conditions in order to increase the efficiency of the main body. Cooling techniques are provided. One such cooling technique is an impingement cooling technique.

FIG. 5 shows an example of a conventional impingement cooling device for a split ring of the gas turbine. In FIG. 5, reference numeral 1 denotes a split ring, that is, an object to be cooled; The cooling medium 5 passes through the inlet hole 3 of the plate 4 and becomes an impingement jet 6, the pressure of which drops to the cavity 10. At the stage of introduction, the object to be cooled 1 is cooled. The above cavity 10
The cooling medium 5 that has cooled the object 1 to be cooled is a small-diameter outlet hole 8.
a flows out into the hot gas 9 from a.

[0004]

In the cooling means utilizing the impingement effect by passing the cooling medium 5 through the impingement cooling holes (entrance holes) 3 as described above, the cooling medium (air, steam, etc.) 5) is generated by the compressor of the gas turbine or supplied from the outside, so that it is required to minimize the consumption of the cooling medium 5 to ensure these efficiencies. Further, in the case of the above-described perforated impingement cooling, a cross flow 7 crossing the impingement jet 6 is generated, and the cooling performance is remarkably reduced due to the cross flow.

Further, when such impingement cooling is applied to a gas turbine or the like, it is usually combined with film cooling. In this case, the high-temperature gas 9 does not flow backward through the film cooling hole 8 in the cavity 10. It is necessary to keep the cavity pressure high. As a result, on the downstream side where the pressure on the high-temperature gas 9 side is low, the differential pressure between the gas surface and the cavity 10 tends to increase, and the cooling medium 5 leaks to the high-temperature gas 9 side, increasing the consumption.

SUMMARY OF THE INVENTION It is an object of the present invention to improve the cooling performance by preventing the occurrence of cross flow across an impingement jet and to reduce the pressure difference between a cavity through which a cooling medium flows and a gas passage to reduce the cooling fluid. It is an object of the present invention to provide an impingement cooling device capable of reducing the amount of outflow to a gas passage and reducing the consumption of a cooling medium.

[0007]

SUMMARY OF THE INVENTION The present invention has been made to solve the above-mentioned problem. The first aspect of the present invention is to provide a high-temperature member such as a component of a gas turbine in which a number of cooling holes are formed. In an impingement cooling device in which a cavity facing the cooling section of the high-temperature member is defined by an impingement plate, the cavity is divided into a plurality of cavities by the impingement plate, and a flow path is restricted between the cavities. The impingement plate is provided with the cooling hole for each of the plurality of cavities.

Further, the second means includes at least the first means.
In the means, the impingement plate is provided with a cooling hole at the inlet portion so that a cooling medium flowing out from the cooling hole to the upstream cavity collides with an upstream cooling surface of the high-temperature member, The cooling hole at the outlet portion is provided such that the cooling medium ejected from the cooling hole through the downstream cavity collides with the downstream cooling surface of the high-temperature member.

According to the above means, the cooling medium is blown out from the impingement cooling hole at the inlet portion of the upstream cavity to the upstream cavity to cool the upstream portion of the high-temperature member and then pass through the gap, and Cooling the downstream portion of the high-temperature member from the impingement cooling hole at the outlet of the impingement cooling by performing the impingement cooling by a plurality of throttles with the same cooling fluid, the cooling medium is:
A sufficient pressure drop is provided to the gas passage inlet, the pressure difference between the downstream cavity and the gas passage is greatly reduced as compared with the conventional one, and the amount of the cooling medium discharged into the gas passage is reduced.

In addition, the cavity is divided into an upstream cavity and a downstream cavity, the passages are communicated with a narrowed gap, and the cooling medium is passed through the upstream impingement cooling holes and the downstream impingement cooling holes. The flow of the cooling medium becomes a smooth flow from the upstream side to the downstream side, so that cross flow can be prevented, and the cooling performance is improved.

[0011]

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS An embodiment of the present invention will be described below in detail with reference to FIGS. FIG. 1 is a sectional view of a main part of a shroud portion (A in FIG. 2) showing a first embodiment of the present invention.
FIG. 2 is a perspective view (bird's eye view) of the shroud portion, FIG. 3 is a perspective view showing a second embodiment, and FIG. 4 is a perspective view showing a third embodiment.

In the first embodiment shown in FIGS.
Reference numeral 17 denotes a gas turbine stationary blade shroud (hereinafter, referred to as a shroud). Reference numeral 17 denotes a partition wall provided at a substantially central portion of the shroud 13 in a width direction (a direction perpendicular to the mainstream gas 15).
Reference numeral 11 denotes a first impingement plate made of a thin heat-resistant metal plate, and the frame portions 13a and 13b of the shroud 13.
Are fixed at both ends by welding. Reference numeral 21 denotes a second impingement plate, one end of which is fixed to the upper end edge of the partition wall 17 by welding, and the other end of which is fixed to the frame portion 13b.

A first cavity 16, which is a space formed between a bottom plate portion 13c of the shroud and a half (left half) of the first impingement plate 11, is provided at an upper portion of the shroud 13, A second cavity 19, which is a space formed between the other half (right half) of the first impingement plate 11 and the upper surface of the second impingement plate 20, is provided.
The space is communicated with a gap 18 in which a flow path is narrowed to some extent. Reference numeral 31 denotes an outlet space formed between the lower surface of the second impingement plate and the bottom plate portion 13c.

The first impingement plate 11 is provided with a number of impingement cooling holes 12 communicating the inlet space of the cooling medium 5 with the first cavity 16. 20 has the second
A number of impingement cooling holes 21 communicating the cavity 19 and the outlet space 31 are formed. Reference numeral 14 denotes a film cooling hole that communicates the first cavity 16 with the passage of the mainstream gas 15, and reference numeral 22 denotes a film cooling hole that communicates the outlet space 31 with the passage of the mainstream gas 15.

In the gas turbine stationary blade shroud cooling apparatus configured as described above, the cooling medium 5 is supplied from a compressor of the gas turbine or an external supply source to impingement cooling holes of the first impingement plate 11. 12 at a flow rate resulting from a suitable pressure differential. The cooling side of the shroud 13, which is the object to be cooled, is cooled by the impingement jet generated by the flow. In a gas turbine having a large thermal load or the like of the shroud 13 to be cooled, the film cooling hole 14 is provided. In this case, the pressure in the first cavity 16 is equal to the pressure of the mainstream gas 15.
Side pressure must be maintained above.

After the cooling medium 5 enters the first cavity 16 and cools the shroud 13 to be cooled,
After an appropriate pressure loss and a rise in the temperature of the cooling medium associated with the above-mentioned cooling and passage through the impingement cooling holes 12, etc., the cooling medium passes through the gap 18 between the first impingement plate 11 and the partition wall 17. , Flows into the second cavity 19.
Thereafter, the cooling medium 5 is supplied to the second impingement plate 20.
Again through the impingement cooling holes 21 at a flow rate resulting from a suitable pressure differential. The cooling side of the shroud 13 as the object to be cooled is cooled again by the impingement jet of the cooling medium 5 generated by the flow. Also,
The cooling medium 5 after passing through the second impingement plate 20 is provided according to the structure and heat load of the shroud 13 which is the object to be cooled. Mainstream gas 1 from film cooling hole 22
It is discharged to the 5 side. However, the pressure of the gas surface from the downstream side of the second cavity 19 to the discharge destination tends to be low in the gas turbine or the like on the downstream side in the axial direction. Passes a plurality of times through the impingement plates 11 and 12 where a large pressure loss occurs, so that the pressure of the cooling medium 5 decreases downstream of the second cavity 19 and the differential pressure from the gas surface pressure decreases. Thus, the amount of leakage of the cooling medium 5 into the mainstream gas 15 can be reduced.

As described above, according to the embodiment shown in FIGS. 1 and 2, the cooling medium 5 is supplied from the first impingement plate 11 → the first cavity 16 → the second cavity 19 → the second cavity 19.
Since the impingement plate 20 → the outlet space 31 and the impingement cooling a plurality of times are performed with the same cooling medium, the consumption of the cooling medium 5 can be significantly reduced as compared with the conventional cooling medium.

In general, in a gas turbine or the like, the mainstream gas 15 on the upstream side in the axial direction has a high pressure. Therefore, the pressure in the first cavity 16 on the upstream side must be maintained at a pressure higher than the pressure on the mainstream gas 15 side. However, in the conventional one,
Due to the above limitation, the amount of the cooling medium due to the blowing of the film from the downstream side film cooling hole 22 was excessively large. However, in the present embodiment, the cavity volume downstream of the second cavity 19 in the design stage has a cavity volume,
Since the pressure difference can be adjusted by the impingement hole diameter, the number, the film cooling hole diameter, the number, and the like, appropriate impingement cooling can be performed by the above adjustment, and the amount of the cooling medium can be suppressed. By adjusting the flow rate of the film blown out from the film cooling hole 22, high film cooling efficiency can be obtained.

Further, in the conventional one, it was difficult to suppress the cross flow which significantly deteriorates the impingement cooling performance. However, according to the embodiment of the present invention, a plurality of impingement plates 11 and 20 are used. , Because the cross flow can be eliminated once each time,
Extremely high impingement cooling performance can be obtained.

In a third embodiment shown in FIG. 3, the present invention is applied to a cooling blade of a gas turbine. In this case,
A first cavity 16 surrounded by an outer surface of a first impingement plate 11 in which a number of impingement cooling holes 12 are drilled and an outer wall 40a of the moving blade, inside a hollow moving blade or a cooling moving blade 40, A second cavity (19) surrounded by a second impingement plate (20) having an impingement cooling hole (21) provided therein; a gap (18) having a narrow flow path between the outlets on both sides of the first cavity (16) and the second cavity (19);
Is communicated with.

In this embodiment, the cooling medium 5 is stepped down from a space surrounded by the first impingement plate 11 through a number of impingement cooling holes 12 and enters the first cavity 16, where two gaps 18 are formed. The pressure is reduced to enter the second cavity 19, the pressure is reduced from the second cavity 19 through a number of impingement cooling holes 21, and flows out into the space between the outer surface of the second impingement plate 20 and the outer wall 40 a. Then, the pressure is sufficiently lowered and flows out from the film cooling hole 22 into the flow path of the mainstream gas 15.

In the third embodiment shown in FIG. 4, the present invention is applied to a turbine split ring of a gas turbine, and the structure in this case is the same as that of the first embodiment shown in FIG.

[0023]

The present invention is configured as described above.
According to the present invention, since impingement cooling is performed by a plurality of throttles with the same cooling medium, the cooling fluid has a sufficient pressure drop to the gas inlet passage inlet, and the differential pressure between the downstream cavity outlet and the gas passage is reduced. Is significantly reduced as compared with the conventional one. Thus, the amount of the cooling medium flowing into the gas passage is reduced, and the consumption of the cooling medium can be reduced.

Further, impingement cooling holes are provided at the inlet and the outlet of the two cavities which are communicated with each other through the narrowed flow path, and the cooling fluid is supplied through the cooling holes and the two cavities in order from the upstream side to the downstream side. By allowing the cooling medium to pass through, the flow of the cooling medium in which the generation of the cross flow is prevented is smooth, and the cooling performance is improved.

[Brief description of the drawings]

FIG. 1 is a cross-sectional view (a cross-sectional view along line AA in FIG. 2) of a main part of an impingement cooling device for a gas turbine vane shroud according to a first embodiment of the present invention.

FIG. 2 is a bird's-eye view of a stationary blade shroud according to the first embodiment.

FIG. 3 is a perspective view of an impingement cooling device for a cooling bucket according to a second embodiment of the present invention.

FIG. 4 is a perspective view of a gas turbine split ring impingement cooling device according to a third embodiment of the present invention.

FIG. 5 is a perspective view showing a conventional gas turbine impingement cooling device.

[Explanation of symbols]

 Reference Signs List 5 cooling medium 11 first impingement plate 12 impingement cooling hole (upstream side) 13 shroud 14 film cooling hole 15 mainstream gas 16 first cavity 18 gap 19 second cavity 20 second impingement plate 21 impingement cooling hole (downstream) Side) 22 Film cooling hole 31 Exit space 40 Cooling blade

 ──────────────────────────────────────────────────続 き Continuing from the front page (72) Inventor Kenichiro Takeishi 2-1-1 Shinama, Arai-machi, Takasago City, Hyogo Prefecture Inside the Takasago Research Laboratory, Mitsubishi Heavy Industries, Ltd. (72) Inventor Kiyoshi Suenaga 2-1-1 Shinama, Araimachi, Takasago City, Hyogo Prefecture No. 1 Inside the Mitsubishi Heavy Industries, Ltd. Takasago Factory

Claims (2)

[Claims] 1. An impingement cooling apparatus comprising: a high-temperature member such as a component of a gas turbine; and an impingement plate having a plurality of cooling holes formed therein to define a cavity facing a cooling portion of the high-temperature member. The cavities are divided into a plurality of cavities by the impingement plate, and a gap is formed between the cavities with a narrow flow path. The impingement plate is provided with the cooling holes for each of the plurality of cavities. An impingement cooling device, comprising: The impingement plate is provided with at least a cooling hole at the inlet so that a cooling medium flowing out of the cooling hole into the upstream cavity collides with an upstream cooling surface of the high-temperature member. 2. The impingement cooling according to claim 1, wherein at least the cooling hole at the outlet portion is provided so that a cooling medium ejected from the cooling hole through the downstream cavity collides with a downstream cooling surface of the high-temperature member. apparatus.