# US 2020/0400068 A1 · Compression in a gas turbine engine

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

- Number: US 2020/0400068 A1
- Document ID: US20200400068A1
- Status: Published application
- Kind: A1
- Application number: 16/558,417
- Filed: 2019-09-03
- 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/US20200400068A1

## CPC

- F02C 3/04: having a turbine driving a compressor (power transmission arrangements F02C7/36; control of working fluid flow F02C9/16)
- F02C 3/107: with two or more rotors connected by power transmission
- 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
- F05D 2200/14: Division
- F05D 2220/323: for aircraft propulsion, e.g. jet engines
- F05D 2220/36: specially adapted for the fan of turbofan engines
- F05D 2240/307: related to the tip of a rotor blade
- F05D 2260/4031: as in toothed gearing
- F05D 2260/40311: of the epicyclical, planetary or differential type
- 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

An engine core including turbine, compressor, and core shaft connecting the turbine to the compressor, wherein a compressor exit temperature has an average airflow; and a fan upstream including a plurality of fan blades extending from a hub, each fan blade having a leading and trailing edge, wherein a fan rotor entry temperature has an average airflow across the leading edge of each blade at cruise conditions and fan tip rotor exit temperature has an average temperature of airflow across a radially outer portion of each blade at the trailing edge cruise conditions. A fan tip temperature rise as: the   fan   tip   rotor   exit   temperature the   fan   rotor   entry   temperature . A core temperature rise as: the   compressor   exit   temperature the   fan   rotor   entry   temperature , A core to fan tip temperature rise ratio of: the   core   temperature   rise the   fan   tip   temperature   rise is in the range from 2.845-3.8.

## Claims

### Claim 1

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 a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor, the engine core further comprising 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 slitter, 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, 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 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 and a fan tip temperature rise is defined as:
the


fan


tip


rotor


exit


temperature

the


fan


rotor


entry


temperature

,
and a core temperature rise is defined as:
the


compressor


exit


temperature

the


fan


rotor


entry


temperature

,
and wherein 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 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 a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor, the engine core having 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 and wherein a 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, 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 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 and a fan tip temperature rise is defined as:
the


fan


tip


rotor


exit


temperature
the


fan


rotor


entry


temperature
,
and a core temperature rise is defined as:
the


compressor


exit


temperature

the


fan


rotor


entry


temperature

,
and wherein 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 3 (depends on 1)

The gas turbine engine of claim 1, wherein the core to fan tip temperature rise ratio is in the range from 2.845 to 3.800.

### Claim 4 (depends on 1)

The gas turbine engine of claim 1, wherein the fan tip temperature rise is in the range from 1.05 to 1.11.

### Claim 5 (depends on 1)

The gas turbine engine of claim 1, wherein the core temperature rise is in the range from 3.1 to 4.0.

### Claim 6 (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, and wherein the radially outer portion of each fan blade is 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 the engine comprises more than one compressor, and wherein the compressor exit temperature is measured at the exit from the highest pressure compressor.

### Claim 8 (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 9 (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 Q has a value in the range from 0.025 to 0.038 Kgs−1N−1K1/2 at cruise conditions.

### Claim 10 (depends on 1)

The gas turbine engine according to claim 1, 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 velocity of the fan tip is in the range from 0.25 to 0.4, at cruise conditions.

### 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, and/or means the conditions experienced by the aircraft and engine at the midpoint between top of climb and start of decent.

### 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 1160 m.

### Claim 16 (depends on 1)

The gas turbine engine according to claim 1, wherein the Overall Pressure Ratio (OPR) at cruise is greater than 40 and lower than 80.

### Claim 17 (depends on 1)

The gas turbine engine 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, and wherein, the gearbox has a gear ratio in the range of from 3.2 to 5.

### 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 core to fan tip temperature rise ratio is in the range from 2.845 to 3.8.