# US 11,111,820 B2 · Gas turbine for aircraft

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- Number: US 11,111,820 B2
- Document ID: US11111820B2
- Status: Granted
- Kind: B2
- Application number: 17/042,446
- Filed: 2018-12-26
- Published: 2021-03-18
- Granted: 2021-09-07
- Earliest claimed: 2018-03-30
- Assignees: Mitsubishi Heavy Industries Aero Engines, Ltd.
- Applicants: Mitsubishi Heavy Industries Aero Engines, Ltd.
- Inventors: Daigo Fujimura; Hiroshi Watanabe; Tadayuki Hanada; Takaaki Hase
- Web page: https://patentgrep.com/patent/US11111820B2

## CPC

- F01D 5/143: {Contour of the outer or inner working fluid flow path wall, i.e. shroud or hub contour}
- F01D 25/24: Casings (modified for heating or cooling F01D25/14); Casing parts, e.g. diaphragms, casing fastenings (casings for rotary machines or engines in general F16M {; special arrangements in stators dealing with breaking-off of part of rotor F01D21/045})
- B64D 27/10: of gas-turbine type (jet aircraft B64D27/16)
- F01D 5/02: Blade-carrying members, e.g. rotors (rotors of non-bladed type F01D1/34; stators F01D9/00 {; selecting particular materials F01D5/28})
- F01D 5/225: {by shrouding}
- F01D 9/02: Nozzles; Nozzle boxes; Stator blades; Guide conduits {, e.g. individual nozzles (nozzle boxes F01D9/047)}
- F01D 9/041: {using blades (F01D5/148 takes precedence)}
- F05D 2220/32: in gas turbines
- F05D 2240/12: Fluid guiding means, e.g. vanes
- F05D 2240/24: for turbines
- F05D 2250/38: angled, e.g. sweep angle
- F01D: NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES (machines or engines for liquids F03; non-positive displacement pumps F04D)
- B64D: EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- F05D: INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS

## Abstract

A gas turbine for an aircraft includes: blades each including a blade body extending outward from a rotor rotating about an axis in a radial direction of the rotor and a fin protruding from an outer peripheral surface of a blade shroud formed on a distal end of the blade body; a casing installed to form gaps between the casing and the fins; and vanes each including a vane shroud formed such that at least a leading end portion of the vane shroud formed upstream in an axial direction is located outside with respect to an extended line of an inner peripheral surface of the casing in the radial direction.

## Claims

### Claim 1

A gas turbine for an aircraft comprising:
a rotor which is capable of rotating about an axis;
blades each including a blade body extending outward from the rotor in a radial direction of the rotor, a blade shroud formed on a distal end of the blade body, and a fin protruding from an outer peripheral surface of the blade shroud;
a casing surrounding the rotor and the blades so as to form gaps between the casing and the fins; and
vanes each including a vane shroud fixed to the casing at a downstream side of the blade in an axial direction of the axis, of which at least a leading end portion formed upstream in the axial direction is located outside with respect to a first extended line of an inner peripheral surface of the casing in the radial direction about the axis, and a vane body extending inward from the vane shroud in the radial direction,
wherein a leading edge of the vane body formed upstream in the axial direction includes: a receding start point; a radial outer end portion which is connected to the vane shroud and is receded downstream in the axial direction with respect to the receding start point; and a leading edge curved portion which is formed to connect the receding start point and the radial outer end portion and which is arched upstream in the axial direction, and
wherein the receding start point is located outside a rear end of an inner surface of the blade shroud in the radial direction.

### Claim 2 (depends on 1)

The gas turbine for the aircraft according to claim 1, wherein the vane shroud includes: a shroud main body portion which is fixed to the casing; and a leading edge fillet portion which is formed between the shroud main body portion and the radial outer end portion of the vane body and is formed so as to be gradually spread toward at least upstream in the axial direction as it approaches the shroud main body portion, and
wherein a distance between the blade shroud and the shroud main body portion in the radial direction is larger than a sum of a length of the leading edge fillet portion in the radial direction and a distance between the blade shroud and the casing in the radial direction.

### Claim 3 (depends on 1)

The gas turbine for the aircraft according to claim 1, wherein the blade shroud is installed at a flare angle so as to be separated from the axis as it approaches downstream in the axial direction,
wherein the casing includes an inclined surface which is formed downstream the gap in the axial direction and which has at an angle corresponding to the flare angle, and
wherein the leading end portion of the vane shroud formed upstream in the axial direction is located outside with respect to a second extended line of the inclined surface in the radial direction.