Gas turbine engine with optimized fan, core passage inlet, and compressor forward stage diameter ratios

1. An aircraft gas turbine engine comprising:

a fan coupled to a fan drive turbine, the fan being configured to provide a bypass flow and a core flow in use;

a reduction gearbox which couples the fan to the fan drive turbine;

a core compressor arrangement, the core compressor arrangement having a core inlet at an upstream end of a core gas flow passage defined by radially inner and outer walls, the radially inner wall of the core inlet defining a first diameter; wherein

the core compressor arrangement comprises an axially forwardmost compressor rotor blade, a root leading edge of the axially forwardmost compressor rotor blade defining a second diameter, wherein a first ratio of the first diameter to the second diameter is greater than or equal to 1.4 and less than or equal to 3; and

wherein a first axial distance is defined by an axial distance between the core inlet and the root leading edge of the axially forwardmost compressor rotor blade, a second ratio is defined by a ratio of the first axial distance divided by a difference between the first diameter and the second diameter, wherein the second ratio is greater than or equal to 0.8, and less than or equal to 1.5.

2. The engine according to claim 1 , wherein the first ratio is 1.5.

3. The engine according to claim 1 , wherein the first ratio is 1.75.

4. The engine according to claim 1 , wherein the second ratio is greater than or equal to 0.8, and less than or equal to 1.2.

5. The engine according to claim 1 , wherein a third ratio is defined by a ratio between a diameter at the leading edge at a root of the fan and a diameter at a leading edge of a tip of the fan, wherein the third ratio is greater than or equal to 0.2 and less than or equal to 0.4.

6. The engine according to claim 5 , wherein the third ratio is 0.25.

7. The engine according to claim 1 , wherein a fourth ratio is defined by a ratio between the diameter at the leading edge of a tip of the fan and the diameter at the leading edge at the root of the first core compressor rotor blade, wherein the fourth ratio is greater than or equal to 3 and less than or equal to 6.

8. The engine according to claim 1 , wherein a fifth ratio is defined by a ratio between the diameter at the leading edge of the tip of the fan and the diameter at the core inlet, wherein the fifth ratio is greater than or equal to 2.5 and less than or equal to 3.5.

9. The engine according to claim 1 , wherein a sixth ratio is defined by a ratio between the diameter at the leading edge of the tip of the fan and the diameter at a leading edge of the tip of the forwardmost compressor rotor blade, wherein the sixth ratio is greater than or equal to 2.5 and 3.5.

10. The engine according to claim 1 , wherein the core compressor arrangement comprises a low pressure compressor upstream of a high pressure compressor.

11. The engine according to claim 10 , wherein the low pressure compressor is directly coupled to the fan drive turbine by a low pressure shaft.

12. The engine according to claim 10 , wherein the low pressure compressor is configured to provide a pressure ratio between 2:1 and 4:1.

13. The engine according to claim 10 , wherein the high pressure compressor is configured to provide a pressure ratio between 10:1 and 30:1.

14. The engine according to claim 10 , wherein the low pressure compressor comprises between 2 and 4 stages, and the high pressure compressor may comprise between 8 and 12 stages.

15. The engine according to claim 1 , wherein the gearbox has a reduction ratio of between 2:1 and 5:1.

16. A method comprising:

providing an aircraft gas turbine engine, which includes:

a fan coupled to a fan drive turbine and configured to provide a bypass flow and a core flow in use,

a reduction gearbox which couples the fan to the fan drive turbine, and

a core compressor arrangement having a core inlet at an upstream end of a core gas flow passage defined by radially inner and outer walls, the radially inner wall of the core inlet defining a first diameter, wherein

the core compressor arrangement comprises an axially forwardmost compressor rotor blade, a root leading edge of the axially forwardmost compressor rotor blade defining a second diameter, wherein a first ratio of the first diameter to the second diameter is greater than or equal to 1.4 and less than or equal to 3, and wherein a first axial distance is defined by an axial distance between the core inlet and the root leading edge of the axially forwardmost compressor rotor blade, a second ratio is defined by a ratio of the first axial distance divided by a difference between the first diameter and the second diameter, wherein the second ratio is greater than or equal to 0.8, and less than or equal to 1.5; and

wherein the aircraft gas turbine engine, during operation, has a bypass ratio greater than or equal to 13 and less than or equal to 25, wherein the bypass ratio is defined by a ratio of air mass flow passing through the fan and around the core inlet, to airflow passing through the core inlet.

17. The method of claim 16 , wherein the engine provides an overall pressure ratio of between 40:1 and 80:1 during operation of the engine.

18. The method of claim 16 , wherein the fan has a fan pressure ratio of between 1.31 and 1.5 during operation of the engine.