Nacelle short inlet

1. A fan assembly for a gas turbine engine, comprising:

a fan including a plurality of fan blades, each fan blade extending radially outwardly from a fan hub to a blade tip, the plurality of blade tips defining a fan diameter;

a nacelle surrounding the fan and defining a fan inlet upstream of the fan, relative to an airflow direction into the fan, the nacelle having a forwardmost edge defining an inlet length from the forwardmost edge to a leading edge of a fan blade of the plurality of fan blades, a ratio of inlet length to fan diameter between 0.20 and 0.45; and

a nacelle inner surface defining a nacelle flowpath, the nacelle flowpath having a convex throat portion and a concave diffusion portion between the throat portion and the leading edge of the fan blade along a bottommost portion of the nacelle defined as a sixty circumferential degree portion of the nacelle centered on a bottom dead center of the nacelle;

wherein a first ratio of inlet length to fan diameter at a bottommost portion of the nacelle is between 0.20 and 0.45, and a second ratio of inlet length to fan diameter at a topmost portion of the nacelle is greater than the first ratio and also between 0.20 and 0.45.

2. The fan assembly of claim 1 , wherein the throat portion has a minimum throat radius measured from a fan central axis less than a maximum diffusion portion radius measured relative to the fan central axis.

3. The fan assembly of claim 1 , wherein the throat portion has a minimum throat radius measured from a fan central axis less than a blade tip radius measured relative to the fan central axis.

4. The fan assembly of claim 1 , wherein a maximum diffusion portion radius measured from a fan central axis is greater than a blade tip radius as measured relative to the central axis.

5. The fan assembly of claim 1 , wherein the fan assembly has a fan pressure ratio between 1.2 and 1.45.

6. The fan assembly of claim 1 , wherein the plurality of fan blades extend radially outwardly from a fan hub at a hub diameter, a ratio of the hub diameter to the fan diameter in the range of 0.25 to 0.45.

7. A gas turbine engine, comprising:

a turbine; and

a fan assembly operably connected to the turbine including:

a fan including a plurality of fan blades, each fan blade extending radially outwardly from a fan hub to a blade tip, the plurality of blade tips defining a fan diameter;

a nacelle surrounding the fan and defining a fan inlet upstream of the fan, relative to an airflow direction into the fan, the nacelle having a forwardmost edge defining an inlet length from the forwardmost edge to a leading edge of a fan blade of the plurality of fan blades, a ratio of inlet length to fan diameter between 0.20 and 0.45; and

a nacelle inner surface defining a nacelle flowpath, the nacelle flowpath having a convex throat portion and a concave diffusion portion between the throat portion and the leading edge of the fan blade along a bottommost portion of the nacelle defined as a sixty circumferential degree portion of the nacelle centered on a bottom dead center of the nacelle;

wherein a first ratio of inlet length to fan diameter at a bottommost portion of the nacelle is between 0.20 and 0.45, and a second ratio of inlet length to fan diameter at a topmost portion of the nacelle is greater than the first ratio and also between 0.20 and 0.45.

8. The gas turbine engine of claim 7 , wherein the throat portion has a minimum throat radius measured from a fan central axis less than a maximum diffusion portion radius measured relative to the fan central axis.

9. The gas turbine engine of claim 7 , wherein the throat portion has a minimum throat radius measured from a fan central axis less than a blade tip radius measured relative to the fan central axis.

10. The gas turbine engine of claim 7 , wherein a maximum diffusion portion radius measured from a fan central axis is greater than a blade tip radius as measured relative to the central axis.

11. The gas turbine engine of claim 7 , wherein the fan assembly has a fan pressure ratio between 1.2 and 1.45.

12. The gas turbine engine of claim 7 , wherein the plurality of fan blades extend radially outwardly from a fan hub at a hub diameter, a ratio of the hub diameter to the fan diameter in the range of 0.25 to 0.45.

13. The gas turbine engine of claim 7 , further comprising a spinner disposed at a central fan axis further defining the nacelle flowpath between the spinner and the nacelle inner surface.

14. A nacelle for a fan of a gas turbine engine, comprising:

a forwardmost edge defining an inlet length from the forwardmost edge to a leading edge of a fan blade of a fan, a ratio of inlet length to a fan diameter between 0.20 and 0.45; and

a nacelle inner surface defining a nacelle flowpath, the nacelle flowpath having a convex throat portion and a concave diffusion portion between the throat portion and the leading edge of the fan blade along a bottommost portion of the nacelle defined as a sixty circumferential degree portion of the nacelle centered on a bottom dead center of the nacelle;

wherein a first ratio of inlet length to fan diameter at a bottommost portion of the nacelle is between 0.20 and 0.45, and a second ratio of inlet length to fan diameter at a topmost portion of the nacelle is greater than the first ratio and also between 0.20 and 0.45.

15. The nacelle of claim 14 , wherein the throat portion has a minimum throat radius measured from a fan central axis less than a maximum diffusion portion radius measured relative to the fan central axis.

16. The nacelle of claim 14 , wherein a maximum diffusion portion radius measured from a fan central axis is greater than a fan radius as measured relative to the central axis.

17. The fan assembly of claim 1 , wherein the ratio of inlet length to fan diameter is 0.20 at a bottommost portion of the nacelle.