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US-12618563-B2 - Lean burn combustion system

US12618563B2US 12618563 B2US12618563 B2US 12618563B2US-12618563-B2

Abstract

A method of operating a gas turbine engine having a lean burn staged combustion system having a combustor in which fuel is combusted. The gas turbine engine includes a fuel-oil heat exchanger arranged to transfer heat between oil and fuel that is provided to the combustor. The method includes transferring heat from the oil to the fuel before the fuel enters the combustor so as to lower the fuel viscosity to 0.58 mm 2 /s or lower on entry to the combustor at cruise conditions.

Inventors

  • Christopher P Madden
  • Craig W Bemment
  • Paul W Ferra
  • Benjamin J KEELER
  • Andrea Minelli
  • Peter SWANN
  • Martin K Yates

Assignees

  • ROLLS-ROYCE PLC

Dates

Publication Date
20260505
Application Date
20230620
Priority Date
20221221

Claims (19)

  1. 1 . A method of operating a gas turbine engine, the gas turbine engine comprising: a lean burn staged combustion system including a combustor in which a fuel is combusted, and a fuel-oil heat exchanger arranged to transfer a heat between an oil and the fuel that is provided to the combustor, the combustor comprising a plurality of fuel spray nozzles including: a first subset of the plurality of fuel spray nozzles that provides a first fuel flow, and a second subset of the plurality of fuel spray nozzles that provides a second fuel flow, wherein the first fuel flow is greater than the second fuel flow below a staging point by having the fuel wholly delivered by pilot injectors of the first subset, and the first subset of fuel spray nozzles is arranged nearer to one or more ignitors of the combustor than the second subset of fuel spray nozzles, the method comprising transferring the heat from the oil to the fuel before the fuel enters the combustor so as to lower a fuel viscosity to 0.58 mm 2 /s or lower on entry to the combustor at cruise conditions, wherein the fuel comprises a sustainable aviation fuel (SAF), and the SAF is present in the fuel in an amount up to and including 100%.
  2. 2 . The method of claim 1 , wherein the heat is transferred from the oil to the fuel before the fuel enters the combustor so as to lower the fuel viscosity to 0.48 mm 2 /s or lower on entry to the combustor at the cruise conditions.
  3. 3 . The method of claim 2 , wherein the heat is transferred from the oil to the fuel before the fuel enters the combustor so as to lower the fuel viscosity to between 0.40 mm 2 /s and 0.48 mm 2 /s on entry to the combustor at the cruise conditions.
  4. 4 . The method of claim 1 , wherein a number of the plurality of fuel nozzles is between 18 and 26 and/or a number of the plurality of fuel spray nozzles per unit engine core size is in a range of 2 to 7.
  5. 5 . The method of claim 4 , wherein each fuel spray nozzle of the plurality of fuel spray nozzles comprises a duplex fuel spray nozzle.
  6. 6 . The method of claim 1 , wherein: the first subset comprises between 2 and 6 fuel spray nozzles.
  7. 7 . The method of claim 1 , wherein the gas turbine engine further comprises a splitter valve arranged to split the fuel flow between the plurality of fuel spray nozzles of the combustor such that the pilot injectors of the first subset receive more the fuel than pilot injectors of the second subset below the staging point.
  8. 8 . The method of claim 1 , further comprising operating the gas turbine engine using the fuel with a lubricity of between 0.71 mm and 0.90 mm wear scar diameter (WSD).
  9. 9 . A gas turbine engine for an aircraft, comprising: a lean burn staged combustion system including a combustor in which a fuel comprising a sustainable aviation fuel (SAF) is combusted, the SAF being present in the fuel in an amount up to and including 100%, the combustor comprising a plurality of fuel spray nozzles including: a first subset of the plurality of fuel spray nozzles that provides a first fuel flow, and a second subset of the plurality of fuel spray nozzles that provides a second fuel flow, wherein the first fuel flow is greater than the second fuel flow below a staging point by having the fuel wholly delivered by pilot injectors of the first subset, and the first subset of fuel spray nozzles is arranged nearer to one or more ignitors of the combustor than the second subset of fuel spray nozzles; a fuel-oil heat exchanger arranged to transfer a heat between an oil and the fuel that is provided to the combustor; and a controller configured to control operation of the fuel-oil heat exchanger to transfer the heat from the oil to the fuel before the fuel enters the combustor so as to lower a fuel viscosity to 0.58 mm 2 /s or lower on entry to the combustor at cruise conditions.
  10. 10 . The gas turbine engine of claim 9 , wherein the controller is configured to control the operation of the fuel-oil heat exchanger to lower the fuel viscosity to 0.48 mm 2 /s or lower on entry to the combustor at the cruise conditions.
  11. 11 . The gas turbine engine of claim 10 , wherein the controller is configured to control the operation of the fuel-oil heat exchanger to lower the fuel viscosity to between 0.40 mm 2 /s and 0.48 mm 2 /s on entry to the combustor at the cruise conditions.
  12. 12 . The gas turbine engine of claim 9 , wherein a number of the plurality of fuel nozzles is between 18 and 26; and/or a number of the plurality of fuel spray nozzles per unit engine core is in a range of 2 to 7.
  13. 13 . The gas turbine engine of claim 12 , wherein each fuel spray nozzle of the plurality of fuel spray nozzles comprises a duplex fuel spray nozzle.
  14. 14 . The gas turbine engine of claim 9 , wherein: the first subset comprises between 2 and 6 fuel spray nozzles.
  15. 15 . The gas turbine engine of claim 9 , further comprises a splitter valve arranged to split the fuel flow between the plurality of fuel spray nozzles of the combustor such that the pilot injectors of the first subset receive more the fuel than pilot injectors of the second subset below the staging point.
  16. 16 . The gas turbine engine of claim 9 , further comprising a fuel distribution system configured to deliver the fuel to the combustor, the fuel having a lubricity of between 0.71 mm and 0.90 mm wear scar diameter (WSD).
  17. 17 . A method of operating a gas turbine engine, the gas turbine engine comprising a lean burn staged combustion system including a combustor in which a fuel is combusted, the combustor comprising a plurality of fuel spray nozzles including: a first subset of the plurality of fuel spray nozzles that provides a first fuel flow, and a second subset of fuel spray nozzles that provides a second fuel flow, wherein the first fuel flow is greater than the second fuel flow below a staging point by having the fuel wholly delivered by pilot injectors of the first subset, and the first subset of fuel spray nozzles is arranged nearer to one or more ignitors of the combustor than the second subset of fuel spray nozzles, the method comprising operating the gas turbine engine using the fuel with a lubricity of between 0.75 mm and 0.90 mm wear scar diameter (WSD), wherein the fuel comprises sustainable aviation fuel (SAF), and the SAF is present in the fuel in an amount up to and including 100%.
  18. 18 . The method according to claim 17 , wherein the SAF is present in the fuel in an amount in a range of 10% to 100%.
  19. 19 . A gas turbine engine for an aircraft, comprising: a lean burn staged combustion system including a combustor in which a fuel comprising a sustainable aviation fuel (SAF) is combusted, the SAF being present in the fuel in an amount up to and including 100%, the combustor comprising a plurality of fuel spray nozzles including: a first subset of the plurality of fuel spray nozzles that provides a first fuel flow, and a second subset of the plurality of fuel spray nozzles that provide a second fuel flow, wherein the first fuel flow is greater than the second fuel flow below a staging point by having the fuel wholly delivered by pilot injectors of the first subset, and the first subset of fuel spray nozzles is arranged nearer to one or more ignitors of the combustor than the second subset of fuel spray nozzles; and a fuel distribution system configured to deliver the fuel to the combustor, the fuel having a lubricity of between 0.75 mm and 0.90 mm wear scar diameter (WSD).

Description

CROSS-REFERENCE TO RELATED APPLICATIONS This specification is based upon and claims the benefit of priority from UK Patent Application Number 2219388.2 filed on 21 Dec. 2022, the entire contents of which are incorporated herein by reference. BACKGROUND Field of the Disclosure The present disclosure relates to a method of operating a gas turbine engine using fuels different from traditional kerosene-based jet fuels. Description of the Related Art There is an expectation in the aviation industry of a trend towards the use of fuels different from the traditional kerosene-based jet fuels generally used at present. SUMMARY According to a first aspect there is provided a method of operating a gas turbine engine, the gas turbine engine having a lean burn staged combustion system in which fuel is combusted. The gas turbine engine comprises a fuel-oil heat exchanger arranged to transfer heat between oil and fuel that is provided to the combustor. The method comprises transferring heat from the oil to the fuel before the fuel enters the combustor so as to lower the fuel viscosity to 0.58 mm2/s or lower on entry to the combustor at cruise conditions. The inventors have appreciated that fuel viscosity has an effect on how the fuel is delivered into and ignited in the combustor (for example, droplet size from fuel spray nozzles, which may impact atomisation and burn efficiency). Taking the fuel viscosity into account when delivering fuel to the combustor, and controlling it as appropriate by varying heat input to the fuel, may therefore provide more efficient fuel burn, improving aircraft performance. A lower viscosity of the fuel at cruise may lend itself to a more efficient engine. The method may comprise transferring heat from the oil to the fuel before the fuel enters the combustor so as to lower the fuel viscosity to 0.48 mm2/s or lower on entry to the combustor at cruise conditions. The method may comprise transferring heat from the oil to the fuel before the fuel enters the combustor so as to lower the fuel viscosity to between 0.50 mm2/s and 0.35 mm2/s, or between 0.48 mm2/s and 0.40 mm2/s, or between 0.44 mm2/s and 0.42 mm2/s on entry to the combustor at cruise conditions. The inventors have determined that a lower bound for the viscosity should take into account the fuel pump operation as too low a fuel viscosity (e.g., as a result of too much heat being put into the fuel from a fuel-oil heat exchanger) could adversely impact the lubrication of bearings within the pump, potentially leading to more wear on the pump, overheating and a failure of the pump. The method may comprise transferring heat from the oil to the fuel before the fuel enters to the combustor so as to lower the fuel viscosity to 0.57, 0.56, 0.55, 0.54, 0.53, 0.52, 0.51, 0.50, 0.49, 0.48, 0.47, 0.46, 0.45, 0.44, 0.43, 0.42, 0.41, 0.40, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31 or 0.30 mm2/s or lower on entry to the combustor at cruise conditions, or a range defined between any two of these values. The method may comprise transferring heat from the oil to the fuel before the fuel enters the combustor so as to raise the fuel temperature to an average of at least 135° C. on entry to the combustor at cruise conditions, or to an average of between 135° C. and 170° C., or to an average of between 135° C. and 160° C., or to an average of between 140° C. and 150° C. on entry to the combustor at cruise conditions. The method may comprise operating the gas turbine engine using a fuel having a lubricity of between 0.71 mm and 0.90 mm wear scar diameter, WSD, or having a lubricity of between 0.75 mm and 0.90 mm WSD, or having a lubricity of between 0.80 mm and 0.90 mm WSD, or having a lubricity of between 0.80 mm and 0.85 mm WSD. According to a second aspect there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising a lean burn staged combustion system having a combustor in which fuel is combusted. The gas turbine engine comprises a fuel-oil heat exchanger arranged to transfer heat between oil and fuel that is provided to the combustor. The gas turbine engine comprises a controller configured to control operation of the fuel-oil heat exchanger to transfer heat from the oil to the fuel before the fuel enters the combustor so as to lower the fuel viscosity to 0.58 mm2/s or lower on entry to the combustor at cruise conditions. The controller may be configured to control operation of the fuel-oil heat exchanger to transfer heat from the oil to the fuel before the fuel enters the combustor so as to lower the fuel viscosity to 0.48 mm2/s or lower on entry to the combustor at cruise conditions. The controller may be configured to control operation of the fuel-oil heat exchanger to transfer heat from the oil to the fuel before the fuel enters the combustor so as to lower the fuel viscosity to between 0.50 mm2/s and 0.35 mm2/s, or between 0.48 mm2/s and 0.40 mm2/s, or between 0.44 mm2/s and 0.42 mm2/s on entry to the combust