# US 11,111,791 B2 · Gas turbine engine having fan diameter ratio

> Agents: this data is available as JSON from the patentgrep API and MCP server — see https://patentgrep.com/llms.txt

- Number: US 11,111,791 B2
- Document ID: US11111791B2
- Status: Granted
- Kind: B2
- Application number: 17/082,886
- Filed: 2020-10-28
- Published: 2021-04-22
- Granted: 2021-09-07
- Earliest claimed: 2019-05-02
- Assignees: ROLLS-ROYCE plc
- Applicants: ROLLS-ROYCE plc
- Inventors: Chathura K Kannangara; Jillian C Gaskell; Stewart T Thornton; Timothy Philp
- Web page: https://patentgrep.com/patent/US11111791B2

## CPC

- F01D 5/14: Form or construction (selecting particular materials, measures against erosion or corrosion F01D5/28)
- B64D 29/06: Attaching of nacelles, fairings or cowlings
- F01D 25/162: {Bearing supports}
- F01D 25/243: {Flange connections; Bolting arrangements (F01D25/265 takes precedence)}
- F01D 25/28: Supporting or mounting arrangements, e.g. for turbine casing
- F02C 7/20: Mounting or supporting of plant; Accommodating heat expansion or creep
- F02C 7/36: Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user ({F02C3/107 - F02C3/13 and} F02C7/32 take precedence; couplings for transmitting rotation F16D; gearing in general F16H)
- F02K 3/06: with front fan
- F16H 1/28: with gears having orbital motion
- F05D 2220/32: in gas turbines
- F05D 2220/36: specially adapted for the fan of turbofan engines
- F05D 2240/14: Casings or housings protecting or supporting assemblies within
- F05D 2240/24: for turbines
- F05D 2240/90: Mounting on supporting structures or systems
- F05D 2260/31: Retaining bolts or nuts
- F05D 2260/40311: of the epicyclical, planetary or differential type
- Y02T 50/60: Efficient propulsion technologies, e.g. for aircraft
- 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
- 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)
- F16H: GEARING
- F05D: INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- Y02T: CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION

## Abstract

A gas turbine engine includes an engine core and a fan located upstream of the engine core. The engine core includes: a compressor system including first, lower pressure compressor, and second, higher pressure compressor; and an outer core casing surrounding the compressor system and including a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection. The first flange connection is the first flange connection that is downstream of an axial position defined by the axial midpoint between the mid-span axial location on the trailing edge of the most downstream aerofoil of the first compressor and the mid-span axial location on the leading edge of the most upstream aerofoil of the second compressor. A fan diameter ratio of: first ⁢ ⁢ flange ⁢ ⁢ radius fan ⁢ ⁢ diameter is equal to or greater than 0.125.

## Claims

### Claim 1

A gas turbine engine for an aircraft comprising:
an engine core comprising:
a compressor system with compressor blades comprising respective aerofoils, the compressor system comprising a first, lower pressure, compressor, and a second, higher pressure, compressor; and
an outer core casing surrounding the compressor system and comprising:
a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection, the first flange connection having a first flange radius, the first flange connection being the first flange connection that is downstream of an axial position defined by an axial midpoint between a mid-span axial location on a trailing edge of a most downstream aerofoil of the first compressor and a mid-span axial location on a leading edge of a most upstream aerofoil of the second compressor; and
a fan located upstream of the engine core, the fan comprising a plurality of fan blades and having a fan diameter between 330 cm and 380 cm;
wherein a fan diameter ratio of:
the
⁢
⁢
first
⁢
⁢
flange
⁢
⁢
radius
the
⁢
⁢
fan
⁢
⁢
diameter
is equal to or greater than 0.125.

### Claim 2 (depends on 1)

The gas turbine engine of claim 1, wherein the fan diameter ratio is less than or equal to 0.17.

### Claim 3 (depends on 1)

The gas turbine engine of claim 1, wherein the fan diameter is between 335 cm and 360 cm.

### Claim 4 (depends on 1)

The gas turbine engine of claim 1, wherein a number of the fan blades is between 16 and 22.

### Claim 5 (depends on 1)

The gas turbine engine of claim 1, further comprising a gearbox that is configured to receive an input from a core shaft and output drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, and optionally wherein a gear ratio of the gearbox is between 3.1 and 4.0.

### Claim 6 (depends on 1)

The gas turbine engine of claim 1, wherein the first flange connection is at, or axially downstream of, the leading edge of the most upstream aerofoil of the second compressor.

### Claim 7 (depends on 1)

The gas turbine engine of claim 1, wherein the first flange connection is at, or axially upstream of, the leading edge of the most upstream aerofoil of the second compressor.

### Claim 8 (depends on 1)

The gas turbine engine of claim 1, further comprising:
a first turbine, and a first core shaft connecting the first turbine to the first compressor; and
a second turbine and a second core shaft connecting the second turbine to the second compressor, and wherein
the second turbine, the second compressor, and the second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.

### Claim 9 (depends on 1)

The gas turbine engine of claim 1, wherein:
the engine core further comprises an inner core casing provided radially inwardly of the compressor blades of the compressor system, the inner core casing and the outer core casing defining a core working gas flow path therebetween,
a gas path radius is defined as an outer radius of the core working gas flow path at the axial position of the first flange connection, and
a gas path ratio of:
the
⁢
⁢
first
⁢
⁢
flange
⁢
⁢
ra
⁢
⁢
dius
the
⁢
⁢
gas
⁢
⁢
path
⁢
⁢
radius
is equal to or greater than 1.10 and less than or equal to 2.0.

### Claim 10 (depends on 1)

The gas turbine engine of claim 1, wherein a fan blade mass ratio of:
the
⁢
⁢
first
⁢
⁢
flange
⁢
⁢
radius
a
⁢
⁢
mass
⁢
⁢
of
⁢
⁢
each
⁢
⁢
fan
⁢
⁢
bl
⁢
⁢
ade
is equal to or less than 19.0 mm/lb.

### Claim 11 (depends on 1)

The gas turbine engine of claim 1, wherein a blade set mass ratio of:
the
⁢
⁢
first
⁢
⁢
⁢
flange
⁢
⁢
radius
a
⁢
⁢
total
⁢
⁢
mass
⁢
⁢
of
⁢
⁢
the
⁢
⁢
plurality
⁢
⁢
of
⁢
⁢
fan
⁢
⁢
blades
is in the range between 0.95 mm/lb and 0.35 mm/lb.

### Claim 12 (depends on 1)

The gas turbine engine of claim 1, further comprising a front mount arranged to be connected to a pylon.

### Claim 13 (depends on 12)

The gas turbine engine of claim 12, wherein a front mount position ratio of:
an
⁢
⁢
axial
⁢
⁢
distance
⁢
⁢
⁢
between
⁢
⁢
⁢
the
⁢
⁢
first
⁢
⁢
⁢
flange
⁢
⁢
connection
⁢
⁢
and
the
⁢
⁢
front
⁢
⁢
mount
⁢
the
⁢
⁢
first
⁢
⁢
flange
⁢
⁢
radius
is equal to or less than 1.18.

### Claim 14 (depends on 12)

The gas turbine engine of claim 12, wherein a front mount position to fan diameter ratio of:
an
⁢
⁢
axial
⁢
⁢
distance
⁢
⁢
between
⁢
⁢
the
⁢
⁢
first
⁢
⁢
flange
⁢
⁢
connection
⁢
⁢
and
the
⁢
⁢
front
⁢
⁢
mount
⁢
the
⁢
⁢
fan
⁢
⁢
diameter
is less than or equal to 0.145.

### Claim 15 (depends on 12)

The gas turbine engine of claim 12, wherein the front mount is a core mount.

### Claim 16 (depends on 1)

The gas turbine engine of claim 1, further comprising:
a nacelle surrounding the engine core and defining a bypass duct between the engine core and the nacelle; and
a fan outlet guide vane (OGV) extending radially across the bypass duct between an outer surface of the engine core and an inner surface of the nacelle, the fan OGV having a radially inner edge and a radially outer edge, wherein an axial midpoint of the radially inner edge is defined as a fan OGV root centrepoint.

### Claim 17 (depends on 16)

The gas turbine engine of claim 16, wherein a fan OGV root position ratio of:
an
⁢
⁢
axial
⁢
⁢
distance
⁢
⁢
between
⁢
⁢
the
⁢
⁢
first
⁢
⁢
flange
⁢
⁢
connection
⁢
⁢
and
the
⁢
⁢
fan
⁢
⁢
OGV
⁢
⁢
root
⁢
⁢
centrepoint
the
⁢
⁢
first
⁢
⁢
flange
⁢
⁢
radius
is equal to or less than 2.6.

### Claim 18 (depends on 16)

The gas turbine engine of claim 16, wherein a fan OGV root position to fan diameter ratio of:
an
⁢
⁢
axial
⁢
⁢
distance
⁢
⁢
between
⁢
⁢
the
⁢
⁢
first
⁢
⁢
flange
⁢
⁢
connection
⁢
⁢
and
the
⁢
⁢
fan
⁢
⁢
OGV
⁢
⁢
root
⁢
⁢
centrepoint
the
⁢
⁢
fan
⁢
⁢
diameter
is less than or equal to 0.33.

### Claim 19 (depends on 16)

The gas turbine engine of claim 16, wherein either or both of (i) and (ii) are satisfied:
(i) a fan OGV tip position ratio of:
an
⁢
⁢
axial
⁢
⁢
distance
⁢
⁢
between
⁢
⁢
the
⁢
⁢
first
⁢
⁢
flange
⁢
⁢
connection
⁢
⁢
and
the
⁢
⁢
fan
⁢
⁢
OGV
⁢
⁢
⁢
tip
⁢
⁢
centrepoint
the
⁢
⁢
first
⁢
⁢
flange
⁢
⁢
radius
is equal to or less than 1.8, and
(ii) a fan OGV tip position to fan diameter ratio of:
the
⁢
⁢
⁢
axial
⁢
⁢
distance
⁢
⁢
between
⁢
⁢
the
⁢
⁢
first
⁢
⁢
flange
⁢
⁢
connection
⁢
⁢
and
the
⁢
⁢
fan
⁢
⁢
OGV
⁢
⁢
⁢
tip
⁢
⁢
centrepoint
the
⁢
⁢
fan
⁢
⁢
diameter
is less than or equal to 0.22.