Combustor size rating for a gas turbine engine using hydrogen fuel
A gas turbine engine includes a hydrogen fuel delivery assembly configured to deliver a hydrogen fuel flow, a compressor section configured to compress air flowing therethrough to provide a compressed air flow, and a combustor including a combustion chamber having a burner length and a burner dome height. The combustion chamber is configured to combust a mixture of the hydrogen fuel flow and the compressed air flow. The combustion chamber can be characterized by a combustor size rating between one inch and seven inches. In more detail, the combustion chamber can be characterized by the combustor size rating between one inch and seven inches at a core air flow parameter between two and one half kN and sixty kN, in which the combustor size rating is a function of the core air flow parameter.
1. A gas turbine engine comprising:
a hydrogen fuel delivery assembly configured to deliver a hydrogen fuel flow;
a compressor section configured to compress air flowing therethrough to provide a compressed air flow; and
a combustion section comprising:
at least one fuel-air mixing assembly comprising a center body, an outer wall spaced from and circumscribing the center body, an annular flow passage defined between the outer wall and the center body, and having an inlet and an outlet, and at least one fuel orifice having a fuel outlet opening into the annular flow passage, and
a combustor configured to operate without diluent, the combustor including an inner liner, an outer liner, and a combustion chamber, the combustion chamber characterized by a combustor size rating between one inch and seven inches at a core air flow parameter between two and one half kN and sixty kN,
wherein the combustor size rating is a function of the core air flow parameter, and
wherein the combustor size rating is defined by:
L
2
H
wherein His a maximum height of the combustion chamber measured by a forward line extending from an inner surface of the outer liner to an inner surface of the inner liner and L is a length of the combustion chamber measured from a midpoint of the forward line to a midpoint of an aft line, the aft line extending from the inner surface of the inner liner to the inner surface of the outer liner at a leading edge of a turbine nozzle,
and, wherein the core air flow parameter is defined by:
Thrust
B
ypass
Ratio
.
2. The gas turbine engine of claim 1 , wherein the center body extends axially from a fore end to an aft end and the inlet is at the fore end and the outlet is at the aft end.
3. The gas turbine engine of claim 2 , wherein the annular flow passage includes a constant cross-sectional area portion along a predetermined portion of the center body and terminating at the aft end.
4. The gas turbine engine of claim 3 , wherein the fuel outlet opening is at a predetermined distance from the aft end of the center body.
5. The gas turbine engine of claim 4 wherein the predetermined distance is less than or equal to 50% of a center body length.
6. The gas turbine engine of claim 4 wherein the predetermined distance is less than 25% of a center body length.
7. The gas turbine engine of claim 6 wherein the fuel outlet opens at the constant cross-sectional area portion.
8. The gas turbine engine of claim 3 wherein the predetermined distance is between 0.0 to 2.0 centimeters from the aft end.
9. The gas turbine engine of claim 3 wherein the fuel outlet opens at the constant cross-sectional area portion.
10. The gas turbine engine of claim 1 wherein the center body comprises a fuel cavity and the at least one fuel orifice has a fuel inlet fluidly coupled to the fuel cavity and wherein the at least one fuel orifice extends through the center body.
11. The gas turbine engine of claim 10 wherein the fuel cavity comprises a channel, extending in an aft-to-fore direction, and the fuel inlet is fluidly coupled to the channel.
12. The gas turbine engine of claim 11 where the channel extends an axial distance between 2% to 50% of a center body length.
13. The gas turbine engine of claim 1 wherein the at least one fuel orifice extends through the outer wall.
14. The gas turbine engine of claim 1 wherein the at least one fuel orifice comprises at least a first set of fuel orifices axially or radially spaced from a second set of fuel orifices.
15. The gas turbine engine of claim 14 wherein the second set of fuel orifices is circumferentially offset from the first set of fuel orifices or at least some of the fuel orifices, of at least one of the first set of fuel orifices or the second set of fuel orifices, are radially angled or axially angled relative to a center body axis.
16. The gas turbine engine of claim 1 wherein the annular flow passage includes a constant cross-sectional area portion along a predetermined portion of the center body and terminating at an aft end and wherein the annular flow passage comprises a reducing cross-sectional area portion located upstream of the constant cross-sectional area portion.
17. The gas turbine engine of claim 16 wherein the reducing cross-sectional area portion terminates at the beginning of the constant cross-sectional area portion.
18. The gas turbine engine of claim 1 , wherein the combustor size rating is between two inches and three and one quarter inches at a core air flow parameter between two and one half kN and fifty kN.
19. The gas turbine engine of claim 1 , wherein the combustor size rating is based on a thrust of the gas turbine engine.