Aircraft operation

Patent metadata

Field	Value
Publication number	US-11859565-B2
Application number	US-202217853333-A
Country	US
Kind code	B2
Filing date	Jun 29, 2022
Priority date	Dec 21, 2021
Publication date	Jan 2, 2024
Grant date	Jan 2, 2024

Abstract

A gas turbine engine includes: a combustor that combust the fuel and having an exit, a combustor exit temperature (T40) is the average temperature of flow and a combustor exit pressure (P40) is the total pressure there; a turbine including a rotor having a leading edge and a trailing edge, and wherein a turbine rotor entry temperature (T41) is an average temperature of flow at the leading edge and a turbine rotor entry pressure (P41) is the total pressure there; and a compressor having an exit, wherein a compressor exit temperature (T30) is the average temperature of flow at the exit from the compressor and a compressor exit pressure (P30) is the total pressure there (all at cruise conditions). A method of determining at least one fuel characteristic includes changing a fuel supplied to the engine; and determining a change in a relationship between T30 or P30, T40 and T41, or of P40 and P41, respectively.

First claim

We claim:

1. A method of determining at least one fuel characteristic of a fuel provided to a gas turbine engine of an aircraft, the gas turbine engine forming part of a propulsion system and comprising:
a combustor arranged to combust the fuel and having an exit, and wherein a combustor exit temperature—T40—is defined as an average temperature of flow at the combustor exit at cruise conditions;
a turbine comprising a rotor having a leading edge and a trailing edge, and wherein a turbine rotor entry temperature—T41—is defined as an average temperature of flow at the leading edge of the rotor of the turbine at the cruise conditions; and
a compressor having an exit, wherein a compressor exit temperature—T30—is defined as an average temperature of flow at the exit from the compressor at the cruise conditions;

wherein the method comprises:
changing the fuel supplied to the gas turbine engine; and
determining the at least one fuel characteristic of the fuel based on a change in at least one of T30, T40, and T41,

wherein the at least one fuel characteristic comprises at least one of:
i. a percentage of sustainable aviation fuel in the fuel; and
ii. an indication that the fuel is a fossil fuel.

2. The method of claim 1 , wherein the determining the at least one fuel characteristic of the fuel is based on a change in a relationship between T30 and one of T40 and T41.

3. The method of claim 2 , wherein the relationship between T30 and the one of T40 and T41 is a difference between T30 and the one of T40 and T41, the difference being indicative of a temperature rise across the combustor.

4. The method of claim 1 , wherein the propulsion system comprises at least one variable inlet guide vane—VIGV, and wherein no change to the VIGV position is made with respect to changing the fuel until after the at least one fuel characteristic of the fuel has been determined.

5. The method of claim 1 , wherein the changing of the fuel supplied to the gas turbine engine is performed at the cruise conditions.

6. The method of claim 1 , wherein the gas turbine engine comprises multiple compressors, and wherein the compressor exit temperature is defined as the temperature at the exit from the highest pressure compressor.

7. The method of claim 1 , wherein the compressor comprises at least one rotor, each of the at least one rotor having a leading edge and a trailing edge, and wherein the compressor exit temperature is defined as the temperature at the axial position of the trailing edge of the rearmost rotor of the compressor.

8. The method of claim 1 , further comprising sensing a response to changing the fuel.

9. A method of determining at least one characteristic of a fuel provided to a gas turbine engine of an aircraft, the gas turbine engine forming part of a propulsion system and comprising:
a combustor arranged to combust the fuel and having an exit, and wherein a combustor exit pressure—P40—is defined as the total pressure at the combustor exit at cruise conditions;
a turbine comprising a rotor having a leading edge and a trailing edge, and wherein a turbine rotor entry pressure—P41—is defined as the total pressure at the leading edge of the rotor of the turbine at the cruise conditions; and
a compressor having an exit, wherein a compressor exit pressure—P30—is defined as the total pressure at the exit from the compressor at the cruise conditions;

wherein the method comprises:
changing the fuel supplied to the gas turbine engine; and
determining the at least one fuel characteristic of the fuel based on a change in at least one of P30, P40 and P41,

wherein the at least one fuel characteristic comprises at least one of:
i. a percentage of sustainable aviation fuel in the fuel; and
ii. an indication that the fuel is a fossil fuel.

10. The method of claim 9 , wherein the determining the at least one fuel characteristic of the fuel is based on a change in a relationship between P30 and one of P40 and P41.

11. The method of claim 9 , wherein:
a combustor exit temperature—T40—is defined as an average temperature of flow at the combustor exit at the cruise conditions;
a turbine rotor entry temperature—T41—is defined as an average temperature of flow at the leading edge of the rotor of the turbine at the cruise conditions; and
a compressor exit temperature—T30—is defined as an average temperature of flow at the exit from the compressor at the cruise conditions;
wherein the method comprises using at least one of T30, T40, and T41 in the determination, in addition to the one or more pressures.

12. The method of claim 9 , wherein the gas turbine engine comprises multiple compressors, and wherein the compressor exit pressure is defined as the pressure at the exit from the highest pressure compressor.

13. The method of claim 9 , wherein the compressor comprises at least one rotor, each rotor having a leading edge and a trailing edge, and wherein the compressor exit pressure is defined as the pressure at the axial position of the trailing edge of the rearmost rotor of the compressor.

14. A propulsion system for an aircraft comprising:
a gas turbine engine comprising:
a combustor arranged to combust a fuel and having an exit, and wherein a combustor exit temperature—T40—is defined as an average temperature of flow at the combustor exit at cruise conditions;
a turbine comprising a rotor having a leading edge and a trailing edge, and wherein a turbine rotor entry temperature—T41—is defined as an average temperature of flow at the leading edge of the rotor of the turbine at the cruise conditions; and
a compressor having an exit, wherein a compressor exit temperature—T30—is defined as an average temperature of flow at the exit from the compressor at the cruise conditions;

a fuel tank arranged to contain the fuel to power the gas turbine engine; and
a processor programmed to:
change the fuel supplied to the gas turbine engine;
receive data corresponding to a change in at least one of T30, T40 and T41; and
determine at least one fuel characteristic of the fuel based on the change in the at least one temperature, the at least one fuel characteristic comprising at least one of:

i. a percentage of sustainable aviation fuel in the fuel; and
ii. an indication that the fuel is a fossil fuel.

15. The propulsion system of claim 14 , wherein the determination is performed based on a change in a relationship between T30 and one of T40 and T41, the relationship between T30 and the one of T40 and T41 being a difference between T30 and the one of T40 and T41.

16. The propulsion system of claim 14 , further comprising at least one sensor arranged to provide the data corresponding to the at least one of T30, T40 and T41 from which the change in the at least one of T30, T40 and T41 is obtained.

17. A propulsion system for an aircraft comprising:
a gas turbine engine comprising:
a combustor arranged to combust a fuel and having an exit, and wherein a combustor exit pressure—P40—is defined as the total pressure at the combustor exit at cruise conditions;
a turbine comprising a rotor having a leading edge and a trailing edge, and wherein a turbine rotor entry pressure—P41—is defined as the total pressure at the leading edge of the rotor of the turbine at the cruise conditions; and
a compressor having an exit, wherein a compressor exit pressure—P30—is defined as the total pressure at the exit from the compressor at the cruise conditions;

a fuel tank arranged to contain the fuel to power the gas turbine engine; and
a processor programmed to:
change the fuel su

Assignees

• Rolls Royce Plc

Inventors

• Keeler Benjamin J
• Beaven David M
• Bemment Craig W
• Ferra Paul W
• Mcnally Kevin R
• Minelli Andrea
• Yates Martin K

Classifications

• F02C9/28 Primary

  Regulating systems responsive to plant or ambient parameters, e.g. temperature, pressure, rotor speed (F02C9/30 - F02C9/38, F02C9/44 take precedence) · CPC title

• F05D2220/323

  for aircraft propulsion, e.g. jet engines · CPC title

• F05D2240/35

  Combustors or associated equipment · CPC title

• F05D2270/303

  Temperature · CPC title

• F05D2270/309

  Rate of change of parameters · CPC title