# US 2020/0400081 A1 · Gas turbine engine transfer efficiency

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- Number: US 2020/0400081 A1
- Document ID: US20200400081A1
- Status: Published application
- Kind: A1
- Application number: 16/545,003
- 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/US20200400081A1

## CPC

- 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)
- F02K 3/06: with front fan
- F01D 11/14: Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing (rotors with blades adjustable in operation F01D7/00)
- F01D 13/003: {with at least two independent shafts, i.e. cross-compound}
- 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})
- F05D 2200/14: Division
- F05D 2260/40311: of the epicyclical, planetary or differential type
- F05D 2270/313: Air 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)
- F01D: NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES (machines or engines for liquids F03; non-positive displacement pumps F04D)
- 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 including a first, lower pressure, turbine, a first compressor, and a first core shaft connecting the first turbine to the first compressor; and a second, higher pressure, turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor, and a fan located upstream of the engine core and including a plurality of fan blades extending from a hub. A low pressure turbine temperature change is defined as: the   first   turbine   exit   temperature the   first   turbine   entrance   temperature . A fan tip temperature rise is defined as: the   fan   tip   rotor   exit   temperature the   fan   rotor   entry   temperature . A turbine to fan tip temperature change ratio of: the   low   pressure   turbine   temperature   change the   fan   tip   temperature   rise is in the range from 1.46 to 2.0.

## Claims

### Claim 1

A gas turbine engine for an aircraft comprising:
an engine core comprising a first turbine, a first compressor, and a first core shaft connecting the first turbine to the first compressor; and a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor, the second turbine being a higher pressure turbine than the first turbine and the second compressor being a higher pressure compressor than the first compressor, wherein a first turbine entrance temperature is defined as an average temperature of airflow at the entrance to the first turbine at cruise conditions and a first turbine exit temperature is defined as an average temperature of airflow at the exit from the first turbine at cruise conditions, and a low pressure turbine temperature change is defined as:
the


first


turbine


exit


temperature
the


first


turbine


entrance


temperature
,
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 splitter, form a streamsurface forming a radially inner boundary of a streamtube that contains all of the bypass flow;
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 the 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 wherein a turbine to fan tip temperature change ratio of:
the


low


pressure


turbine


temperature


change
the


fan


tip


temperature


rise
is in the range from 1.46 to 2.0.

### Claim 2

A gas turbine engine for an aircraft comprising:
an engine core having a core radius defined between the centreline of the engine and a forwardmost tip of the engine core, the engine core comprising a first turbine, a first compressor, and a first core shaft connecting the first turbine to the first compressor; and a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor, the second turbine being a higher pressure turbine than the first turbine and the second compressor being a higher pressure compressor than the first compressor, wherein a first turbine entrance temperature is defined as an average temperature of airflow at the entrance to the first turbine at cruise conditions and a first turbine exit temperature is defined as an average temperature of airflow at the exit from the first turbine at cruise conditions, and a low pressure turbine temperature change is defined as:
the


first


turbine


exit


temperature
the


first


turbine


entrance


temperature
;
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 defined as 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 the 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 wherein a turbine to fan tip temperature change ratio of:
the


low


pressure


turbine


temperature


change
the


fan


tip


temperature


rise
is in the range from 1.46 to 2.0.

### Claim 3 (depends on 1)

The gas turbine engine of claim 1, wherein the turbine to fan tip temperature change ratio is in the range from 1.5 to 1.8.

### 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 1.1.

### Claim 5 (depends on 1)

The gas turbine engine of claim 1, wherein the low pressure turbine temperature change is in the range from 1.6 to 1.85.

### Claim 6 (depends on 1)

The gas turbine engine according to claim 1, wherein the engine comprises more than two turbines and wherein the highest pressure turbine of the engine is selected as the second turbine and the lowest pressure turbine of the engine is selected as the first turbine.

### 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−1s.

### 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 (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 (translational) velocity of the fan tip is in the range from 0.25 to 0.4.

### 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 the Overall Pressure Ratio (OPR) at cruise is greater than 40 and lower than 80.

### 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 first core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the first core shaft, and wherein, optionally, 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 first (low pressure) turbine has four or more rotor stages.

### 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 turbine to fan tip temperature change ratio is in the range from 1.46 to 2.0.