Aircraft fuelling
A method of operating a gas turbine engine; the gas turbine engine includes a combustor. The combustor includes a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber. The plurality of fuel spray nozzles includes a first subset of fuel spray nozzles and a second subset of fuel spray nozzles. The combustor is operable in a condition in which the first subset of fuel spray nozzles are supplied with more fuel than the second subset of fuel spray nozzles. A ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5. The method includes: providing fuel to the one or more fuel-oil heat exchangers. Also provided is a gas turbine engine for an aircraft.
1 . A method of operating a gas turbine engine, the gas turbine engine comprising:
a combustor, comprising an annular combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which the first subset of fuel spray nozzles are supplied with more fuel than the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5;
one or more fuel-oil heat exchangers; and
a plurality of ignitors;
the method comprising:
providing a fuel to the one or more fuel-oil heat exchangers;
transferring heat from oil to the fuel in the one or more fuel-oil heat exchangers; and
providing the fuel from the one or more fuel-oil heat exchangers to the plurality of fuel spray nozzles; wherein
heat is transferred from the oil to the fuel in the one or more fuel-oil heat exchangers to raise a temperature of the fuel to an average of between 135° C. and 200° C. on injection of the fuel into the combustion chamber at cruise conditions,
the plurality of fuel spray nozzles are arranged circumferentially spaced around the combustor within a liner of the annular combustion chamber and positioned at a same radial distance from a rotational axis of the gas turbine engine, and
the plurality of fuel spray nozzles are controlled differently such that the first subset of fuel spray nozzles are supplied with more fuel than the second subset of fuel spray nozzles below a staging point,
wherein two of the fuel spray nozzles of the first subset are positioned circumferentially spaced from each other around the combustor,
wherein each of the first subset of fuel spray nozzles is located nearer a respective one or more of the ignitors than the second subset, and
wherein the plurality of ignitors are arranged circumferentially spaced around the combustor.
2 . The method of claim 1 , wherein the plurality of ignitors include 3, 5, 7 or 9 ignitors.
3 . The method of claim 1 , wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:3 to 1:4.
4 . The method of claim 1 , wherein the first subset of fuel spray nozzles includes between 1 and 10 fuel spray nozzles.
5 . The method of claim 1 , wherein the second subset of fuel spray nozzles includes between 10 and 25 fuel spray nozzles.
6 . The method of claim 1 , wherein the fuel provided to the plurality of fuel spray nozzles may comprise a % SAF in the range of 50-100%.
7 . A gas turbine engine for an aircraft, comprising:
a combustor, comprising an annular combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which the first subset of fuel spray nozzles are supplied with more fuel than the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5;
a plurality of ignitors; and
a fuel distribution system configured to provide a fuel to the plurality of fuel spray nozzles, comprising one or more fuel-oil heat exchangers and a controller; wherein
the controller is configured to control the one or more fuel-oil heat exchangers so as to raise a temperature of the fuel to an average of between 135° C. and 200° C. on injection of the fuel into the combustion chamber at cruise conditions, and
the plurality of fuel spray nozzles are arranged circumferentially spaced around the combustor within a liner of the annular combustion chamber and positioned at a same radial distance from a rotational axis of the gas turbine engine, and
the plurality of fuel spray nozzles are controlled differently such that the first subset of fuel spray nozzles are supplied with more fuel than the second subset of fuel spray nozzles below a staging point,
wherein two of the fuel spray nozzles of the first subset are positioned circumferentially spaced from each other around the combustor,
wherein each of the first subset of fuel spray nozzles is located nearer a respective one or more of the ignitors than the second subset, and
wherein the plurality of ignitors are arranged circumferentially spaced around the combustor.
8 . The gas turbine engine of claim 7 , wherein the plurality of ignitors include 3, 5, 7 or 9 ignitors.
9 . The gas turbine engine of claim 7 , wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:3 to 1:4.
10 . The gas turbine engine of claim 7 , wherein the first subset of fuel spray nozzles includes between 1 and 10 fuel spray nozzles.
11 . The gas turbine engine of claim 7 , wherein the second subset of fuel spray nozzles includes between 10 and 25 fuel spray nozzles.
12 . The gas turbine engine of claim 7 , wherein the fuel provided to the plurality of fuel spray nozzles may comprise a % SAF in the range of 50-100%.
13 . The method of claim 1 , wherein
the staging point defines the point at which the combustor switches between pilot-only operation and pilot-and-main operation.
14 . The gas turbine engine of claim 7 , wherein
the staging point defines the point at which the combustor switches between pilot-only operation and pilot-and-main operation.