Cooling structure for recovery-type air-cooled gas turbine combustor

• US 9,512,781 B2
• Filed: 08/10/2011
• Issued: 12/06/2016
• Est. Priority Date: 09/30/2010
• Status: Active Grant

• Current Assignee: Mitsubishi Power Limited (Mitsubishi Heavy Industries, Ltd.), Tohoku Electric Power Company Incorporated
• Original Assignee: mitsubishi hitachi power systems, ltd. (Mitsubishi Heavy Industries, Ltd.), Tohoku Electric Power Company Incorporated
• Inventors: Mizukami, Satoshi, Ishiguro, Tatsuo, Masada, Junichiro, Takata, Kazumasa, Torii, Shunsuke, Fukunaga, Yuya, Konishi, Tetsu, Nishimura, Yoshiaki, Igarashi, Yoshiaki, Ishikawa, Shinya
• Primary Examiner(s): Weiss, Nicholas J
• Assistant Examiner(s): Nguyen, Ngoc T
• Application Number:
  US13/806,523
• Publication Number:
  US 20130098063A1
• Time in Patent Office: 1,945 Days
• US Class Current: 1/1
• CPC Class Codes: F01D 9/023 Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
• CPC Class Codes: F02C 7/08 Heating air supply before combustion, e.g. by exhaust gases
• CPC Class Codes: F02C 7/18 the medium being gaseous, e.g. air F02C7/125 takes precedence
• CPC Class Codes: F02C 7/24 Heat or noise insulation air intakes having provisions for noise suppression F02C7/045; turbine exhaust heads, chambers, or the like F01D25/30; silencing nozzles of jet-propulsion plants F02K1/00
• CPC Class Codes: F05D 2260/202 by film cooling
• CPC Class Codes: F23R 2900/00014 Reducing thermo-acoustic vibrations by passive means, e.g. by Helmholtz resonators
• CPC Class Codes: F23R 3/005 Combined with pressure or heat exchangers
• CPC Class Codes: F23R 3/06 Arrangement of apertures along the flame tube
• CPC Class Codes: F23R 3/44 Combustion chambers comprising a single tubular flame tube within a tubular casing reverse-flow combustion chambers F23R3/54

Abstract

In a cooling structure for a recovery-type air-cooled gas turbine combustor having a recovery-type air-cooling structure that bleeds, upstream of the combustor, and pressurizes compressed air supplied from a compressor, that uses the bled and pressurized air to cool a wall, and that recovers and reuses the bled and pressurized air as combustion air for burning fuel in the combustor together with a main flow of the compressed air, wall cooling in which cooling air is supplied to cooling air passages formed in the wall of the combustor to perform cooling involves a downstream wall region, closer to a turbine, that is cooled using the bled and pressurized air as the cooling air and an upstream wall region, closer to a burner, that is cooled using, as the cooling air, bled compressed air bled from a main flow of the compressed air through a housing inner space.

Claims

1. A cooling structure for a recovery-type air-cooled gas turbine combustor comprising:

first cooling passages formed in a downstream wall region of a wall of the gas turbine combustor, the first cooling passages being configured to cool the downstream wall region of the gas turbine combustor by receiving first cooling air bled from a compressed air supply passage and pressurized; and

second cooling passages formed in an upstream wall region of the wall of the gas turbine combustor, the second cooling passages being configured to cool the upstream wall region of the gas turbine combustor by receiving second cooling air from an inner space of a housing in which the gas turbine is installed, the second cooling air being supplied to the inner space of the housing from the compressed air supply passage,

wherein the first cooling air passages are configured so that the first cooling air flows into the first cooling passages and exits into the inner space of the housing,

wherein the second cooling air passages are configured so that the second cooling air is introduced from the inner space of the housing into the second cooling passages, and

wherein the first cooling air introduced into the first cooling passages for cooling the downstream wall region of the gas turbine combustor is supplied to the gas turbine combustor as combustion air for burning fuel.

2. The cooling structure for the gas turbine combustor according to claim 1, wherein:

the first cooling passages include first air inlet holes and first air outlet holes;

the first air inlet holes are provided closer to the turbine relative to the first air outlet holes, which are provided closer to a burner of the gas turbine combustor;

in the first cooling passages, the first cooling air introduced into the first air inlet holes provided closer to the turbine flows out of the first air outlet holes provided closer to the burner in the downstream wall region into the inner space of the housing; and

in the second cooling passages, at least the second cooling air introduced from second air inlet holes provided closer to the turbine in the upstream wall region flows out of second air outlet holes provided in a center of the upstream wall region.

3. The cooling structure for the gas turbine combustor according to claim 2, wherein the first air outlet holes and the second air inlet holes that are adjacent to each other near a boundary between the upstream wall region and the downstream wall region are arranged in a staggered pattern.

4. The cooling structure for the gas turbine combustor according to claim 2, wherein at least the first cooling passages that direct the first cooling air introduced into the first air inlet holes to the first air outlet holes include turn-back portions that couple together the first cooling passages that are adjacent to each other in a circumferential direction of the gas turbine combustor near the boundary.

5. The cooling structure for the gas turbine combustor according to claim 4, wherein the turn-back portions are routed so as to cover a region near the boundary to perform wall cooling using the first cooling air and the second cooling air.

6. The cooling structure for the gas turbine combustor according to claim 4, wherein the second cooling passages, which direct the second cooling air introduced from the second air inlet holes to the second air outlet holes, include turn-back portions that couple together the second cooling passages that are adjacent to each other in a circumferential direction of the gas turbine combustor near the boundary.

7. The cooling structure for the gas turbine combustor according to claim 2, wherein a partition wall protruding from the wall of the combustor is provided between the first air outlet holes and the second air inlet holes.

8. The cooling structure for the gas turbine combustor according to claim 2, wherein a ratio (P/d) of a pitch (P) of the first air outlet holes and the second air inlet holes that are opened and adjacent to each other near the boundary in a longitudinal direction of the combustor to the diameter (d) of the first air outlet holes is set to 2 or more (P/d≧

2).

9. The cooling structure for the gas turbine combustor according to claim 2, wherein a diameter (di) of the second air inlet holes is set to a value larger than the diameter (d) of the first air outlet holes (di>

d).

10. The cooling structure for the gas turbine combustor according to claim 1, wherein a boundary between the upstream wall region and the downstream wall region is located near an end of an acoustic liner closer to the turbine.

11. The cooling structure for the gas turbine combustor according to claim 1, wherein at least the second cooling passages that direct the second cooling air introduced from the second air inlet holes to the second air outlet holes include turn-back portions that couple together the second cooling passages that are adjacent to each other in a circumferential direction of the gas turbine combustor near the boundary.