Propulsion system arrangement for turbofan gas turbine engine

1. An integrated propulsion system comprising components that include a gas turbine engine, a nacelle assembly and a mounting assembly, the system designed by a process comprising identifying two or more of internal structural loading requirements, external structural mount loading requirements, aerodynamic requirements, and acoustic requirements for the system, and interdependently designing said components to meet said requirements, wherein:

the nacelle assembly includes a fan nacelle and an aft nacelle, the fan nacelle arranged at least partially about a fan and the engine, and the fan nacelle arranged at least partially about a core cowling to define a bypass flow path;

the system includes a fan section at least partially defined by the fan nacelle, and the fan section has a low fan pressure ratio of less than 1.7;

the engine includes a turbine section driving a compressor section, the turbine section includes a fan drive turbine and a high pressure turbine, the compressor section includes a low pressure compressor and a high pressure compressor, and a shaft interconnecting the fan and the fan drive turbine;

the fan section is configured to deliver a portion of air into the compressor section, and a portion of air into a bypass duct defining the bypass flow path; and

the aft nacelle is arranged at least partially about the core cowling to define an aft flowpath portion of the bypass flow path, the aft flowpath portion extending circumferentially between opposite sides of a single bifurcation positioned in the bypass duct such that the aft flowpath portion is uninterrupted by another bifurcation between the opposite sides.

2. The integrated propulsion system as recited in claim 1 , comprising:

a pylon attaching the nacelle assembly and the engine to an aircraft structure.

3. The integrated propulsion system as recited in claim 2 , comprising a thrust reverser that selectively communicates air from the bypass duct.

4. The integrated propulsion system as recited in claim 3 , wherein the pylon comprises the mounting assembly, the mounting assembly attachable to the nacelle assembly and the engine, the mounting assembly including a forward mount and an aft mount, the forward mount attachable to one of a fan case surrounding the fan and an intermediate case, and the aft mount attachable to one of a mid-turbine frame and a turbine exhaust case.

5. The integrated propulsion system as recited in claim 3 , further comprising a geared architecture, and wherein the shaft is connected to the fan through the geared architecture to drive the fan at a lower speed than the fan drive turbine, and the geared architecture is an epicyclic gear train.

6. The integrated propulsion system as recited in claim 5 , wherein the aft nacelle is moveable along an engine longitudinal axis relative to the fan nacelle.

7. The integrated propulsion system as recited in claim 6 , wherein:

the fan drive turbine has a first exit area at a first exit point and is rotatable at a first speed, the high pressure turbine has a second exit area at a second exit point and is rotatable at a second speed, and the second speed being faster than the first speed; and

a first performance quantity is defined as the product of the first speed squared and the first area, a second performance quantity is defined as the product of the second speed squared and the second area, and a performance ratio of the first performance quantity to the second performance quantity is between 0.5 and 1.5.

8. The integrated propulsion system as recited in claim 7 , wherein the low fan pressure ratio is less than 1.45.

9. The integrated propulsion system as recited in claim 8 , wherein the performance ratio is greater than or equal to 1.0.

10. The integrated propulsion system as recited in claim 3 , wherein the pylon comprises the mounting assembly, the mounting assembly attachable to the nacelle assembly and the engine, the mounting assembly including a forward mount and an aft mount, the forward mount is attachable to a fan case, and the aft mount is attachable to a mid-turbine frame.

11. The integrated propulsion system as recited in claim 10 , wherein the mid-turbine frame supports a bearing system, and the mid-turbine frame includes airfoils in a core flow path.

12. The integrated propulsion system as recited in claim 3 , wherein the pylon comprises the mounting assembly, the mounting assembly attachable to the nacelle assembly and the engine, the mounting assembly including a forward mount and an aft mount, the forward mount is attachable to one of a fan case surrounding the fan and an intermediate case, and the aft mount is attachable to one of a mid-turbine frame and a turbine exhaust case.

13. The integrated propulsion system as recited in claim 12 , wherein the forward mount is attachable to the intermediate case, and the aft mount is attachable to the turbine exhaust case.

14. The integrated propulsion system as recited in claim 12 , wherein the aft nacelle is moveable along an engine longitudinal axis relative to the fan nacelle.

15. The integrated propulsion system as recited in claim 14 , wherein the forward mount is attachable to the fan case.

16. The integrated propulsion system as recited in claim 15 , wherein the aft mount is attachable to the turbine exhaust case.

17. The integrated propulsion system as recited in claim 16 , wherein the mid-turbine frame supports a bearing system, and the mid-turbine frame includes airfoils in a core flow path.

18. The integrated propulsion system as recited in claim 17 , wherein the low fan pressure ratio is less than 1.5.

19. The integrated propulsion system as recited in claim 18 , further comprising a plurality of circumferentially spaced struts extending in the bypass flow path radially between the intermediate case and the fan case, the plurality of struts axially between the fan and the single bifurcation with respect to the engine longitudinal axis.

20. The integrated propulsion system as recited in claim 19 , wherein:

the forward mount reacts to at least a vertical load, a first side load and a torsional load relative to the engine longitudinal axis in operation; and

the aft mount reacts to at least a vertical load and a second side load relative to the engine longitudinal axis in operation.

21. The integrated propulsion system as recited in claim 20 , wherein the mounting assembly includes an intermediate mount attachable to the intermediate case, and the intermediate mount reacts to at least a thrust load relative to the engine longitudinal axis in operation.

22. The integrated propulsion system as recited in claim 21 , wherein the forward mount is attachable to the intermediate mount.

23. The integrated propulsion system as recited in claim 21 , wherein the intermediate mount is attachable to the aft mount.

24. The integrated propulsion system as recited in claim 23 , wherein only the forward mount relative to the mounting assembly reacts to torsional loads in operation.

25. The integrated propulsion system as recited in claim 23 , wherein only the aft mount relative to the mounting assembly reacts to torsional loads in operation.

26. The integrated propulsion system as recited in claim 18 , wherein:

the fan drive turbine has a first exit area at a first exit point and is rotatable at a first speed, the high pressure turbine has a second exit area at a second exit point and is rotatable at a second speed, and the second speed being faster than the first speed; and

a first performance quantity is defined as the product of the first speed squared and the first area, a second performance quantity is defined as the product of the second speed squared and the second area, and a performance ratio of the first performance quantity to the second performance quantity is between 0.2 and 0.4.

27. The integrated propulsion system of claim 26 , wherein:

the aft flowpath portion circumferentially defines an angle between the opposite sides of the single bifurcation, and the angle is greater than or equal to 320 degrees; and

the nacelle assembly is a slim-line nacelle, the slim-line nacelle defines a maximum diameter and has an inlet lip defining a highlight diameter, and a ratio of the highlight diameter to the maximum diameter is greater than or equal to 0.80.