# US 2020/0400080 A1 · Gas turbine engine with highly efficient fan

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- Number: US 2020/0400080 A1
- Document ID: US20200400080A1
- Status: Published application
- Kind: A1
- Application number: 16/545,129
- Filed: 2019-08-20
- Published: 2020-12-24
- Earliest claimed: 2019-06-24
- Assignees: ROLLS-ROYCE plc
- Applicants: Rolls-Royce PLC
- Inventors: Craig W. BEMMENT; Pascal DUNNING
- Web page: https://patentgrep.com/patent/US20200400080A1

## CPC

- F02C 9/20: by throttling; by adjusting vanes
- F02K 3/06: with front fan
- F02C 9/28: Regulating systems responsive to plant or ambient parameters, e.g. temperature, pressure, rotor speed (F02C9/30 - F02C9/38, F02C9/44 take precedence)
- F05D 2220/323: for aircraft propulsion, e.g. jet engines
- F05D 2220/36: specially adapted for the fan of turbofan engines
- F05D 2270/07: to improve fuel economy
- F05D 2270/303: Temperature
- Y02T 50/60: Efficient propulsion technologies, e.g. for 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)
- 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 for an aircraft includes an engine core with a turbine, compressor, and core shaft connecting the two; and a fan upstream of the core with a plurality of blades extending from a hub each with a leading and trailing edge, wherein fan tip radius is between the engine centreline and each blade's leading edge outermost tip and hub radius is between the engine centreline and the hub's outer surface at each blade's leading edge radial position, the ratio of hub to tip radius between 0.2 and 0.285. A fan rotor entry temperature is the average temperature of airflow across the leading edge of each blade at cruise conditions and a fan rotor exit temperature is an average temperature of airflow across a radially outer portion of each blade at the trailing edge at cruise conditions, the ratio of entry to exit temperature between 1.11 and 1.05.

## Claims

### Claim 1

A gas turbine engine for an aircraft comprising:
an engine core comprising a turbine, a compressor, a core shaft connecting the turbine to the compressor, and an annular splitter at which the flow is divided between a core flow that flows through the engine core, and a bypass flow that flows along a bypass duct, wherein stagnation streamlines around the circumference of the engine, stagnating on a leading edge of the annular splitter, form a streamsurface forming a radially inner boundary of a streamtube that contains all of the bypass flow; and
a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, wherein a fan tip radius of the fan is defined between a centreline of the engine and an outermost tip of each fan blade at its leading edge and a hub radius is defined between the centreline of the engine and an outer surface of the hub at a blade root at the radial position of the leading edge of each fan blade, each fan blade having a radially outer portion lying within the streamtube that contains the bypass flow, and wherein a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan rotor exit temperature is defined as an average temperature of airflow across a radially outer portion of each fan blade at the trailing edge at cruise conditions;
wherein a fan hub to tip ratio of:
the


fan


hub


radius
the


fan


tip


radius
is in the range from 0.2 to 0.285; and
a fan tip temperature rise of:
the


fan


tip


rotor


exit


temperature
the


fan


rotor


entry


temperature
is in the range from 1.11 to 1.05, and wherein
a fan tip loading at cruise conditions is defined as dH/Utip2, where dH is the enthalpy rise across the fan and Utip is the (translational) velocity of the fan tip, and the fan tip loading is in the range from 0.25 to 0.4.

### Claim 2

A gas turbine engine for an aircraft comprising:
an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, wherein the engine core has a core radius defined between the centreline of the engine and a forwardmost tip of the engine core; and
a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, wherein a fan tip radius of the fan is defined between a centreline of the engine and an outermost tip of each fan blade at its leading edge and a hub radius is defined between the centreline of the engine and an outer surface of the hub at a blade root at the radial position of the leading edge of each fan blade, a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan tip rotor exit temperature is defined as an average temperature of airflow across a radially outer portion of each fan blade at the trailing edge at cruise conditions, wherein the radially outer portion of each fan blade is or comprises the portion of each fan blade at a distance from the centreline of the engine greater than the core radius;
wherein a fan hub to tip ratio of:
the


fan


hub


radius
the


fan


tip


radius
is in the range from 0.2 to 0.285; and
a fan tip temperature rise of:
the


fan


tip


rotor


exit


temperature
the


fan


rotor


entry


temperature
is in the range from 1.11 to 1.05, and wherein
a fan tip loading at cruise conditions is defined as dH/Utip2, where dH is the enthalpy rise across the fan and Utip is the (translational) velocity of the fan tip, and the fan tip loading is in the range from 0.25 to 0.4.

### Claim 3

(canceled)

### Claim 4 (depends on 1)

The gas turbine engine according to claim 1, wherein the fan tip temperature rise is equal to 1.1.

### Claim 5 (depends on 1)

The gas turbine engine according to claim 1, further comprising a nacelle surrounding the fan and the engine core and defining a bypass duct outside of the engine core, and wherein the fan tip rotor exit temperature and the fan rotor entry temperature each provide a temperature of airflow across the fan blade portion in a bypass stream of air about to enter the bypass duct.

### Claim 6 (depends on 5)

The gas turbine engine according to claim 5 wherein the radially outer portion of each fan blade is or comprises the portion of each fan blade extending across the entrance to the bypass duct.

### Claim 7 (depends on 1)

The gas turbine engine according to claim 1, wherein a specific thrust of the engine at cruise conditions, defined as net engine thrust divided by mass flow rate through the engine, is in the range from 50 to 100 Nkg−1 s.

### Claim 8 (depends on 1)

The gas turbine engine according to claim 1, wherein a quasi-non-dimensional mass flow rate Q is defined as:
Q
=
W

T
0
P
0
·
A
f

a

n
where:
W is mass flow rate through the fan in Kg/s;
T0 is average stagnation temperature of the air at the fan face in Kelvin;
P0 is average stagnation pressure of the air at the fan face in Pa;
Afan is the area of the fan face in m2;
and has a value in the range from 0.025 to 0.038 Kgs−1N−1K1/2 at cruise conditions.

### Claim 9 (depends on 8)

The gas turbine engine according to claim 8, wherein Q takes a value less than or equal to 0.035 Kgs−1N−1K1/2 at cruise conditions.

### Claim 10

(canceled)

### Claim 11 (depends on 1)

The gas turbine engine according to claim 1, wherein cruise conditions means the conditions at mid-cruise of an aircraft to which the engine is attached.

### Claim 12 (depends on 1)

A gas turbine engine according to claim 1, wherein the forward speed of the gas turbine engine at the cruise conditions is in the range of from Mn 0.75 to Mn 0.85.

### Claim 13 (depends on 1)

A gas turbine engine according to claim 1, wherein the cruise conditions correspond to atmospheric conditions defined by the International Standard Atmosphere at an altitude of 11582 m and a forward Mach Number of 0.8.

### Claim 14 (depends on 1)

A gas turbine engine according to claim 1, wherein the cruise conditions correspond to atmospheric conditions defined by the International Standard Atmosphere at an altitude of 10668 m and a forward Mach Number of 0.85.

### Claim 15 (depends on 1)

The gas turbine engine according to claim 1, wherein the cruise conditions correspond to atmospheric conditions at an altitude that is in the range of from 10500 m to 11600 m.

### Claim 16 (depends on 1)

The gas turbine engine according to claim 1, wherein a fan pressure ratio is defined as the ratio of the mean total pressure of the air flow at the exit of the fan to the mean total pressure of the air flow at the inlet of the fan, and wherein, at cruise conditions:
the fan pressure ratio is in a range between 1.2 and 1.45, or
the fan pressure ratio is in a range between 1.35 and 1.43, or
the fan pressure ratio is 1.39.

### Claim 17 (depends on 1)

The gas turbine engine of claim 1, wherein:
(i) the fan tip radius is in the range from 110 cm to 150 cm; or
(ii) the fan tip radius is in the range from 155 cm to 200 cm.

### Claim 18 (depends on 1)

The gas turbine engine according to claim 1, further comprising a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.

### Claim 19 (depends on 1)

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

### Claim 20 (depends on 1)

A method of operating a gas turbine engine on an aircraft, the gas turbine engine being as defined in claim 1, wherein the method comprises:
operating the gas turbine engine to provide propulsion under cruise conditions such that the fan hub to tip ratio is in the range from 0.2 to 0.285, and the fan tip temperature rise is in the range from 1.11 to 1.05.

### Claim 21 (depends on 1)

The gas turbine engine according to claim 1, wherein
a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor,
a core temperature rise is defined as:
the


compressor


exit


temperature
the


fan


rotor


entry


temperature
and a core to fan tip temperature rise ratio of:
the


core


temperature


rise
the


fan


tip


temperature


rise
is in the range from 2.845 to 3.8.

### Claim 22 (depends on 2)

The gas turbine engine according to claim 2, wherein
a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor,
a core temperature rise is defined as:
the


compressor


exit


temperature
the


fan


rotor


entry


temperature
and a core to fan tip temperature rise ratio of:
the


core


temperature


rise
the


fan


tip


temperature


rise
is in the range from 2.845 to 3.8.