Aircraft engine
An aircraft engine comprising a fan, the fan having a diameter D and including a plurality of fan blades, the fan blades having a sweep metric S, each fan blade having a leading edge, and a forward-most portion on the leading edge of each fan blade being in a first reference plane. The aircraft engine further comprises a nacelle, comprising an intake portion forward of the fan, a forward edge on the intake portion being in a second reference plane, wherein the intake portion has a length L measured along an axis of the aircraft engine between the first reference plane and the second reference plane, the aircraft engine having a cruise design point condition M rel , wherein M rel is between 0.4 and 0.93, and L/D is between 0.2 and 0.45.
1. An aircraft engine comprising:
a fan, the fan having a diameter D, the fan including a plurality of fan blades, each fan blade having a sweep metric S and a leading edge, a forward-most portion on the leading edge of each fan blade being in a first reference plane; and
a nacelle, comprising an intake portion upstream of the fan, a forward-most edge on the intake portion being in a second reference plane, wherein the intake portion has a length L measured along an axis of the aircraft engine between the first reference plane and the second reference plane;
wherein the aircraft engine has a relative Mach number at a cruise design point condition M rel from from 0.4 to 0.93 and L/D is from 0.2 to 0.45, and
wherein the sweep metric S of each fan blade is from 0.2 to 0.9.
2. The aircraft engine of claim 1 , wherein the sweep metric S of each fan blade is from 0.25 to 0.9.
3. The aircraft engine of claim 1 , wherein the sweep metric S of each fan blade is from 0.30 to 0.9.
4. The aircraft engine of claim 1 , wherein the sweep metric S of each fan blade is from 0.35 to 0.9.
5. The aircraft engine of claim 1 , wherein the sweep metric S of each fan blade is from 0.40 to 0.9.
6. The aircraft engine of claim 1 , wherein L/D is from 0.25 to 0.40.
7. The aircraft engine of claim 1 , wherein L/D is from 0.30 to 0.40.
8. The aircraft engine of claim 1 , wherein the aircraft engine is a gas turbine engine.
9. The aircraft engine of claim 8 , wherein the gas turbine engine is a geared gas turbine engine.
10. The aircraft engine of claim 9 , wherein the aircraft engine accommodates a first rate of air flow into an intermediate pressure compressor and a second rate of air flow which passes through a bypass duct, a ratio of the second rate of air flow to the first rate of air flow being greater than 10.
11. The aircraft engine of claim 8 , wherein the aircraft engine accommodates a first rate of air flow into an intermediate pressure compressor and a second rate of air flow which passes through a bypass duct, a ratio of the second rate of air flow to the first rate of air flow being greater than 10.
12. The aircraft engine of claim 1 , wherein the aircraft engine is powered by electricity.
13. The aircraft engine of claim 1 , wherein the fan diameter D is from 220 cm to 380 cm.
14. An aircraft comprising at least one aircraft engine according to claim 1 .
15. An aircraft engine comprising:
a fan, the fan having a diameter D, the fan including a plurality of fan blades, each fan blade having a sweep metric S and a leading edge, a forward-most portion on the leading edge of each fan blade being in a first reference plane; and
a nacelle, comprising an intake portion upstream of the fan, a forward-most edge on the intake portion being in a second reference plane, wherein the intake portion has a length L measured along an axis of the aircraft engine between the first reference plane and the second reference plane;
wherein the aircraft engine has a relative Mach number at a cruise design point condition M rel from 0.4 to 0.93 and L/D is from 0.2 to 0.45, and
wherein the sweep metric S of each fan blade is from 0.25 to 0.9.
16. An aircraft engine comprising:
a fan, the fan having a diameter D, the fan including a plurality of fan blades, each fan blade having a sweep metric S and a leading edge, a forward-most portion on the leading edge of each fan blade being in a first reference plane; and
a nacelle, comprising an intake portion upstream of the fan, a forward-most edge on the intake portion being in a second reference plane, wherein the intake portion has a length L measured along an axis of the aircraft engine between the first reference plane and the second reference plane;
wherein the aircraft engine has a relative Mach number at a cruise design point condition M rel from 0.4 to 0.93 and L/D is from 0.2 to 0.45, and
wherein the sweep metric S of each fan blade is from 0.30 to 0.9.