# US 11,994,075 B2 · Geared gas turbine engine

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- Number: US 11,994,075 B2
- Document ID: US11994075B2
- Status: Granted
- Kind: B2
- Application number: 18/135,526
- Filed: 2023-04-17
- Published: 2023-08-31
- Granted: 2024-05-28
- Earliest claimed: 2019-05-23
- Assignees: ROLLS-ROYCE plc
- Applicants: Rolls-Royce PLC
- Inventors: Craig W Bemment
- Web page: https://patentgrep.com/patent/US11994075B2

## CPC

- F02C 3/107: with two or more rotors connected by power transmission
- F02C 9/24: Control of the pressure level in closed cycles
- F01D 19/00: Starting of machines or engines; Regulating, controlling, or safety means in connection therewith (warming-up before starting F01D25/10; turning or inching gear F01D25/34)
- F02C 9/18: by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages {(F02C3/113 takes precedence)}
- F02K 3/06: with front fan
- F02C 7/268: Starting drives for the rotor {, acting directly on the rotor of the gas turbine to be started}
- 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)
- 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)
- F01D: NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES (machines or engines for liquids F03; non-positive displacement pumps F04D)
- 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)
- Y02T: CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION

## Abstract

A gas turbine engine for an aircraft and a method of operating a gas turbine engine on an aircraft. Embodiments disclosed include a gas turbine engine for an aircraft including: an engine core has a turbine, a compressor, and a core shaft; a fan located upstream of the engine core, the fan has a plurality of fan blades; a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle; and a gearbox that receives an input from the core shaft and outputs drive to the fan wherein the gas turbine engine is configured such that a jet velocity ratio of a first jet velocity exiting from the bypass exhaust nozzle to a second jet velocity exiting from an exhaust nozzle of the engine core at idle conditions is greater by a factor of 2 or more than the jet velocity ratio at maximum take-off conditions.

## Claims

### Claim 1

A method of operating a gas turbine engine on an aircraft, the gas turbine engine comprising:
an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor;
a fan located upstream of the engine core, the fan comprising a plurality of fan blades;
a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle; and
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,
wherein the method comprises operating the gas turbine engine to provide propulsion such that a jet velocity ratio, RJ, of a first jet velocity exiting from the bypass exhaust nozzle to a second jet velocity exiting from an exhaust nozzle of the engine core is defined as:
R
J
=
V
B
⁢
C
B
V
C
⁢
C
C
⁢
η
LPT
⁢
η
F
where VB is a fully expanded first jet velocity, CB is a thrust coefficient of the bypass exhaust nozzle, VC is a fully expanded second jet velocity, CC is a thrust coefficient of the engine core exhaust nozzle, ηLPT is an isentropic efficiency of a lowest pressure turbine of the engine core and ηF is an isentropic efficiency of a fan tip; and
the jet velocity ratio, RJ, is between around 0.75 and 1.3 at maximum take-off conditions.

### Claim 2 (depends on 1)

The method of claim 1, wherein the jet velocity ratio RJ, is between around 0.8 and 1.0 at the maximum take-off conditions.

### Claim 3 (depends on 1)

The method of claim 1, wherein the temperature of the flow at the exit of the combustor, at a position immediately upstream of a first turbine vane, is at least 1800K at the maximum take-off conditions.

### Claim 4 (depends on 1)

The method of claim 1,
wherein a temperature of a flow at an exit of the combustor, at a position immediately upstream of a first turbine vane, is at least 1850K at the maximum take-off conditions.

### Claim 5 (depends on 1)

The method of claim 1,
wherein a temperature of a flow at an exit of the combustor, at a position immediately upstream of a first turbine vane, is at least 1900K at the maximum take-off conditions.

### Claim 6 (depends on 1)

The method of claim 1,
wherein a temperature of a flow at an exit of the combustor, at a position immediately upstream of a first turbine vane, is at least 1950K at the maximum take-off conditions.

### Claim 7 (depends on 4)

The method of claim 4, wherein the temperature of the flow at the exit of the combustor, at the position immediately upstream of the first turbine vane, is less than 2000K at the maximum take-off conditions.

### Claim 8 (depends on 1)

The method of claim 1, wherein the jet velocity ratio, RJ, is between around 2 and 3 at idle conditions.

### Claim 9 (depends on 1)

The method of claim 1, wherein the jet velocity ratio, RJ, is between around 0.75 and 1.3 at cruise conditions.

### Claim 10 (depends on 1)

The method of claim 1,
wherein a gear ratio of the gearbox is between 3.1 and 3.3; a fan tip loading defined as dH/Utip2 is between 0.28 and 0.31 at cruise conditions, where dH is an enthalpy rise across the fan and Utip is a translational velocity of the leading edge of the fan tip; and a bypass ratio, defined as a ratio of a mass flow rate of a flow through the bypass duct to a mass flow rate of a flow through the core at cruise conditions, is in a range of from 11.5 to 13.5.

### Claim 11 (depends on 10)

The method of claim 10, wherein an overall pressure ratio defined as a ratio of a stagnation pressure upstream of the fan to a stagnation pressure at an exit of a highest pressure compressor is between 40 and 55 at the cruise conditions.

### Claim 12 (depends on 1)

The method of claim 1, wherein the fan has an outer diameter of around 220 cm.

### Claim 13

A method of operating a gas turbine engine on an aircraft, the gas turbine engine comprising:
an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor;
a fan located upstream of the engine core, the fan comprising a plurality of fan blades;
a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle; and
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,
wherein the method comprises operating the gas turbine engine to provide propulsion such that a jet velocity ratio, RJ, of a first jet velocity exiting from the bypass exhaust nozzle to a second jet velocity exiting from an exhaust nozzle of the engine core is defined as:
R
J
=
V
B
⁢
C
B
V
C
⁢
C
C
⁢
η
LPT
⁢
η
F
where VB is a fully expanded first jet velocity, CB is a thrust coefficient of the bypass exhaust nozzle, VC is a fully expanded second jet velocity, CC is a thrust coefficient of the engine core exhaust nozzle, ηLPT is an isentropic efficiency of a lowest pressure turbine of the engine core and ηF is an isentropic efficiency of a fan tip; and
the jet velocity ratio, RJ, is between around 2 and 3 at idle conditions.

### Claim 14 (depends on 13)

The method of claim 13, wherein the jet velocity ratio, RJ, is between around 0.75 and 1.3 at cruise conditions.

### Claim 15 (depends on 13)

The method of claim 13,
wherein a fan tip loading defined as dH/Utip2 is between 0.28 and 0.35 at cruise conditions, where dH is an enthalpy rise across the fan and Utip is a translational velocity of a leading edge of the fan tip.

### Claim 16 (depends on 13)

The method of claim 13, wherein a specific thrust, defined as a net thrust of the engine divided by a total mass flow through the engine, is between 80 Nkg−1s and 95 Nkg−1s at cruise conditions.

### Claim 17 (depends on 13)

The method of claim 13,
wherein a temperature of a flow at an exit of the combustor, at a position immediately upstream of a first turbine vane, is at least 1600K at cruise conditions.

### Claim 18 (depends on 13)

The method of claim 13,
wherein a temperature of a flow at an exit of the combustor, at a position immediately upstream of a first turbine vane, is between 1600K and 1650K at cruise conditions.

### Claim 19 (depends on 13)

The method of claim 13, wherein the fan comprises 22 or 24 fan blades.

### Claim 20 (depends on 13)

The method of claim 13,
wherein a gear ratio of the gearbox is between 3.1 and 3.4; a fan tip loading defined as dH/Utip2 is between 0.28 and 0.32 at cruise conditions, where dH is an enthalpy rise across the fan and Utip is a translational velocity of a leading edge of the fan tip; a bypass ratio, defined as a ratio of a mass flow rate of a flow through the bypass duct to a mass flow rate of a flow through the core at cruise conditions, is in a range of from 11.5 to 14; an overall pressure ratio defined as a ratio of a stagnation pressure upstream of the fan to a stagnation pressure at an exit of a highest pressure compressor is between 40 and 55 at the cruise conditions; a ratio of a radius of a fan blade of the fan at a hub to a radius of the fan blade at the tip is in a range of 0.27 to 0.31; and a specific thrust, defined as a net thrust of the engine divided by a total mass flow through the engine, is between 80 s and 95 s at cruise conditions.