# US 10,479,519 B2 · Nacelle short inlet for fan blade removal

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- Number: US 10,479,519 B2
- Document ID: US10479519B2
- Status: Granted
- Kind: B2
- Application number: 15/378,300
- Filed: 2016-12-14
- Published: 2017-07-06
- Granted: 2019-11-19
- Earliest claimed: 2015-12-31
- Assignees: United Technologies Corporation
- Applicants: United Technologies Corporation
- Inventors: Yuan J. Qiu; Steven H. Zysman; Robert E. Malecki; Wesley K. Lord
- Web page: https://patentgrep.com/patent/US10479519B2

## CPC

- B64D 33/02: of combustion air intakes
- B64D 27/10: of gas-turbine type (jet aircraft B64D27/16)
- B64D 29/00: Power-plant nacelles, fairings or cowlings
- 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})
- F02C 7/04: Air intakes for gas-turbine plants or jet-propulsion plants
- F02K 3/06: with front fan
- B64D 2033/0273: {for jet engines}
- B64D 2033/0286: {for turbofan engines}
- F05D 2220/32: in gas turbines
- F05D 2220/36: specially adapted for the fan of turbofan engines
- F05D 2230/70: Disassembly methods
- F05D 2240/303: related to the leading edge of a rotor blade
- F05D 2250/51: Inlet
- F05D 2250/711: convex
- B64D: EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- F01D: NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES (machines or engines for liquids F03; non-positive displacement pumps F04D)
- F02C: GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS (construction of turbines F01D; jet-propulsion plants F02K; construction of compressors or fans F04; generating combustion products of high pressure or high velocity F23R; using gas turbines in compression refrigeration plants F25B11/00)
- F02K: JET-PROPULSION PLANTS (features of jet-propulsion plants common to gas-turbine plants, air intakes or fuel supply control of air-breathing jet-propulsion plants F02C7/00, F02C9/00)
- F05D: INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS

## Abstract

A fan assembly for a gas turbine engine includes a fan including a plurality of fan blades extending radially outwardly from a fan hub to a blade tip defining a fan diameter. A nacelle surrounds the fan and defines a fan inlet upstream of the fan, relative to an airflow direction into the fan. The nacelle has a forwardmost edge defining an inlet length from the forwardmost edge to a leading edge of a fan blade. A ratio of inlet length to fan diameter is between 0.20 and 0.45. A nacelle inner surface defines a nacelle flowpath having a convex throat portion and a diffusion portion between the throat portion and the leading edge of the fan blade at a topmost portion of the nacelle. The throat portion has a minimum throat radius measured from a fan central axis greater than a blade tip radius.

## Claims

### Claim 1

A fan assembly for a gas turbine engine, comprising:
a fan including a plurality of fan blades, each fan blade extending radially outwardly from a fan hub to a blade tip, the plurality of blade tips defining a fan diameter;
a nacelle surrounding the fan and defining a fan inlet upstream of the fan, relative to an airflow direction into the fan, the nacelle having a forwardmost edge defining an inlet length from the forwardmost edge to a leading edge of a fan blade of the plurality of fan blades, a ratio of inlet length to fan diameter between 0.20 and 0.45; and
a nacelle inner surface defining a nacelle flowpath, the nacelle flowpath having a convex throat portion and a diffusion portion between the throat portion and the leading edge of the fan blade at a topmost portion of the nacelle, the throat portion having a minimum throat radius measured from a fan central axis greater than a blade tip radius measured relative to the fan central axis.

### Claim 2 (depends on 1)

The fan assembly of claim 1, wherein a ratio of a fan hub outer radius to an inner surface radius measured from the central axis at an axial position corresponding to a forwardmost fan blade tip at the topmost portion of the nacelle is greater than the ratio of fan hub outer radius to inner surface radius at a remainder of a nacelle perimeter.

### Claim 3 (depends on 2)

The fan assembly of claim 2, wherein the ratio of fan hub outer radius to inner surface radius at the topmost portion is in the range of 0.35 to 0.40.

### Claim 4 (depends on 1)

The fan assembly of claim 1, wherein the topmost portion is defined as a sixty circumferential degree portion of the nacelle centered on a top dead center of the nacelle.

### Claim 5 (depends on 1)

The fan assembly of claim 1, wherein the fan assembly has a fan pressure ratio between 1.2 and 1.45.

### Claim 6 (depends on 1)

The fan assembly of claim 1, wherein the plurality of fan blades extend radially outwardly from a fan hub at a hub diameter, a ratio of the hub diameter to the fan diameter in the range of 0.25 to 0.45.

### Claim 7 (depends on 1)

The fan assembly of claim 1, further comprising a spinner disposed at a central fan axis further defining the nacelle flowpath between the spinner and the nacelle inner surface.

### Claim 8 (depends on 7)

The fan assembly of claim 7, wherein a minimum distance between the spinner and the nacelle inner surface is greater than a fan blade radial length.

### Claim 9 (depends on 1)

The fan assembly of claim 1, wherein one or more fan blades are removable via the fan inlet without removal of the nacelle.

### Claim 10

A gas turbine engine, comprising:
a turbine; and
a fan assembly operably connected to the turbine including:
a fan including a plurality of fan blades, each fan blade extending radially outwardly from a fan hub to a blade tip, the plurality of blade tips defining a fan diameter;
a nacelle surrounding the fan and defining a fan inlet upstream of the fan, relative to an airflow direction into the fan, the nacelle having a forwardmost edge defining an inlet length from the forwardmost edge to a leading edge of a fan blade of the plurality of fan blades, a ratio of inlet length to fan diameter between 0.20 and 0.45; and
a nacelle inner surface defining a nacelle flowpath, the nacelle flowpath having a convex throat portion and a diffusion portion between the throat portion and the leading edge of the fan blade at a topmost portion of the nacelle, the throat portion having a minimum throat radius measured from a fan central axis greater than a blade tip radius measured relative to the fan central axis.

### Claim 11 (depends on 10)

The gas turbine engine of claim 10, wherein a ratio of a fan hub outer radius to an inner surface radius measured from the central axis at an axial position corresponding to a forwardmost fan blade tip at the topmost portion of the nacelle is greater than the ratio of fan hub outer radius to inner surface radius at a remainder of a nacelle perimeter.

### Claim 12 (depends on 11)

The gas turbine engine of claim 11, wherein the ratio of fan hub outer radius to inner surface radius at the topmost portion is in the range of 0.35 to 0.40.

### Claim 13 (depends on 10)

The gas turbine engine of claim 10, wherein the topmost portion is defined as a sixty circumferential degree portion of the nacelle centered on a top dead center of the nacelle.

### Claim 14 (depends on 10)

The gas turbine engine of claim 10, wherein the fan assembly has a fan pressure ratio between 1.2 and 1.45.

### Claim 15 (depends on 10)

The gas turbine engine of claim 10, wherein the plurality of fan blades extend radially outwardly from a fan hub at a hub diameter, a ratio of the hub diameter to the fan diameter in the range of 0.25 to 0.45.

### Claim 16 (depends on 10)

The gas turbine engine of claim 10, further comprising a spinner disposed at a central fan axis further defining the nacelle flowpath between the spinner and the nacelle inner surface.

### Claim 17 (depends on 16)

The gas turbine engine of claim 16, wherein a minimum distance between the spinner and the nacelle inner surface is greater than a fan blade radial length.

### Claim 18 (depends on 10)

The gas turbine engine of claim 10, wherein one or more fan blades are removable via the fan inlet without removal of the nacelle.

### Claim 19

A nacelle for a fan of a gas turbine engine, comprising:
a forwardmost edge defining an inlet length from the forwardmost edge to a leading edge of a fan blade of a fan, a ratio of inlet length to a fan diameter between 0.20 and 0.45; and
a nacelle inner surface defining a nacelle flowpath, the nacelle flowpath having a convex throat portion and a diffusion portion between the throat portion and the leading edge of the fan blade at a topmost portion of the nacelle, the throat portion having a minimum throat radius measured from a fan central axis greater than a blade tip radius of a fan blade measured relative to the fan central axis.