Aircraft gas turbine

Patent metadata

Field	Value
Publication number	US-9945250-B2
Application number	US-201113515978-A
Country	US
Kind code	B2
Filing date	Jan 17, 2011
Priority date	Feb 24, 2010
Publication date	Apr 17, 2018
Grant date	Apr 17, 2018

Abstract

An aircraft gas turbine is constituted by accommodating a compressor ( 14 ), a combustor ( 15 ), and a turbine ( 16 ) in a cylindrical main unit casing ( 12 ). A thick wall part ( 52 ) is provided on an outer periphery side of rotor blades ( 34 ) in the main unit casing. A cooling passage ( 53 ), for cooling the thick wall part ( 52 ) by circulating compressed air compressed by the compressor, is provided in the thick wall part. Also provided is a discharge passage ( 55 ) for discharging compressed air having circulated in the cooling passage to a combustion gas passage A. The structural strength of the thick wall part is ensured by appropriately cooling the thick wall part of the casing, while simplifying the structure and preventing a decrease in efficiency, thereby ensuring effective containment performance and an appropriate clearance between the casing and the rotor blades.

First claim

The invention claimed is:

1. An aircraft gas turbine comprising:
a casing constituting an outermost shell in a cylindrical shape to accommodate a compressor, a combustor, and a turbine therein, and including a wall part provided on an outer periphery side of a rotor blade of the turbine;
a shroud disposed radially inside the wall part of the casing so as to form an empty space between the shroud and the wall part of the casing;
a cooling passage provided in the wall part to circulate compressed air compressed by the compressor to cool the wall part; and
a discharge passage for discharging compressed air having circulated in the cooling passage to a combustion gas passage,
wherein the cooling passage includes an inlet passage with an inlet passage first end being open to a suction part of the combustor, an outlet passage with an outlet passage second end being open to the discharge passage, and a return passage with a return passage first end being in communication with an inlet passage second end, and a return passage second end being in communication with an outlet passage first end,
the inlet passage, the return passage, and the outlet passage being separately disposed so as to align in a circumferential direction of the casing at an equal interval such that the cooling passage forms a serpentine shape along the circumferential direction,
the cooling passage is formed in the wall part of the casing,
the empty space formed between the shroud and the wall part of the casing is the discharge passage for discharging compressed air having circulated in the cooling passage to the combustion gas passage,
the discharge passage is in communication with the combustion gas passage via an opening formed in the shroud, and
the discharge passage discharges compressed air having circulated in the cooling passage to an upstream side of a nozzle in the combustion gas passage via the opening formed in the shroud.

2. The aircraft gas turbine according to claim 1 , wherein the cooling passage is provided in the wall part of the casing along an axis direction of a turbine rotor, and is arranged in plural in a circumferential direction at an equal interval.

Assignees

• Kitamura Tsuyoshi
• Masaki Shinichiro
• Mitsubishi Heavy Ind Aero Engines Ltd

Inventors

• Kitamura Tsuyoshi
• Masaki Shinichiro

Classifications

• F02C7/18

  the medium being gaseous, e.g. air {(F02C7/125 takes precedence)} · CPC title

• F01D11/24

  by selectively cooling-heating stator or rotor components · CPC title

• F05D2220/323

  for aircraft propulsion, e.g. jet engines · CPC title

• F05D2240/11

  Shroud seal segments · CPC title

• F01D25/14 Primary

  Casings modified therefor (double casings F01D25/26) · CPC title