Front section stiffness ratio

1. A gas turbine engine comprising:

a fan rotor driven by a fan drive turbine about an axis through a gear reduction to reduce a speed of said fan rotor relative to a speed of said fan drive turbine;

a fan case surrounding said fan rotor, and an inner core housing surrounding a compressor section, including a low pressure compressor;

said fan rotor delivering air into a bypass duct defined between said fan case and a splitter wall of said inner core housing, and a rigid connection between said fan case and said inner core housing including a plurality of A-frames each including a pair of legs rigidly connected at a connection point to said fan case, and each leg in said pair extending away from said connection point in opposed circumferential directions to be connected to a compressor wall of said inner core housing to form an A-shape; and

the rigid connection also including a plurality of fan exit guide vanes rigidly connected to said fan case; and

a lateral stiffness is defined in a radial direction, a lateral stiffness ratio of the lateral stiffness of the plurality of fan exit guide vanes and a lateral stiffness of a combination of the plurality of A-frame, the compressor wall, and a fan intermediate case which is forward of the low pressure compressor being greater than or equal to 0.6 and less than or equal to 2.0.

2. The gas turbine engine as set forth in claim 1 , wherein a moment stiffness is defined about an axis extending perpendicular to a rotational axis of the gas turbine engine, and a moment stiffness ratio of the moment stiffness of the plurality of fan exit guide vanes to a moment stiffness of the combination of the plurality of A-frames, the compressor wall, and the fan intermediate case being greater than or equal to 0.50 and less than or equal to 2.0.

3. The gas turbine engine as set forth in claim 2 , wherein the moment stiffness ratio being less than or equal to 1.4.

4. The gas turbine engine as set forth in claim 2 , wherein the moment stiffness ratio being less than or equal to 1.0.

5. The gas turbine engine as set forth in claim 2 , wherein a first gap is defined between a leading edge of the fan exit guide vanes at a radially innermost point radially outward of the splitter wall and a leading edge of the plurality of A-frames at a radially inner location radially inward of the splitter wall, and a second gap is defined between a trailing edge of the fan blades and a leading edge of the plurality of fan exit guide vanes at the radially inward point radially outward of the splitter wall each of the gaps being defined by extending a line from the respective points in a radially outward direction, and then measuring a distance with a line extending parallel to the rotational axis of the engine, and a first gap ratio of said second gap to said first gap is between 0.25 and 0.45.

6. The gas turbine engine as set forth in claim 5 , wherein a third gap is defined between the trailing edge of the fan blades and a leading edge of the fan exit guide vane at a radially outermost location radially inward of the fan case, the third gap also defined by extending a line from the respective points in a radially outward direction, and then measuring a distance with a line extending parallel to the rotational axis of the engine, and a second gap ratio of said first gap to said third gap is greater than or equal to 1.1 and less than or equal to 2.

7. The gas turbine engine as set forth in claim 6 , wherein a third gap ratio of said second gap to said third gap is greater than or equal to 0.4 and less than or equal to 0.6.

8. The gas turbine engine as set forth in claim 2 , wherein the lateral stiffness ratio being less than or equal to 1.4.

9. The gas turbine engine as set forth in claim 8 , wherein the lateral stiffness ratio being less than or equal to 1.0.

10. The gas turbine engine as set forth in claim 1 , wherein a ratio of a quantity of the plurality of fan exit guide vanes to a quantity of the legs of the plurality of A-frames being greater than or equal to 1.75.

11. The gas turbine engine as set forth in claim 1 , wherein said low pressure compressor has between four to six stages.

12. The gas turbine engine as set forth in claim 1 , wherein a first gap is defined between a leading edge of the fan exit guide vanes at a radially innermost point radially outward of the splitter wall and a leading edge of the plurality of A-frames at a radially inner location radially inward of the splitter wall, and a second gap is defined between a trailing edge of the fan blades and a leading edge of the plurality of fan exit guide vanes at the radially inward point radially outward of the splitter wall each of the gaps being defined by extending a line from the respective points in a radially outward direction, and then measuring a distance with a line extending parallel to the rotational axis of the engine, and a first gap ratio of said second gap to said first gap is between 0.25 and 0.45.

13. The gas turbine engine as set forth in claim 11 , wherein a third gap is defined between the trailing edge of the fan blades and a leading edge of the fan exit guide vane at a radially outermost location radially inward of the fan case, the third gap also defined by extending a line from the respective points in a radially outward direction, and then measuring a distance with a line extending parallel to the rotational axis of the engine, and a second gap ratio of said first gap to said third gap is greater than or equal to 1.1 and less than or equal to 2.

14. The gas turbine engine as set forth in claim 13 , wherein a third gap ratio of said second gap to said third gap is greater than or equal to 0.4 and less than or equal to 0.6.

15. The gas turbine engine as set forth in claim 1 , wherein the lateral stiffness ratio being less than or equal to 1.4.

16. The gas turbine engine as set forth in claim 15 , wherein the lateral stiffness ratio being less than or equal to 1.0.

17. A gas turbine engine comprising:

a fan rotor driven by a fan drive turbine about an axis through a gear reduction to reduce a speed of said fan rotor relative to a speed of said fan drive turbine;

a fan case surrounding said fan rotor, and an inner core housing surrounding a compressor section, including a low pressure compressor;

said fan rotor delivering air into a bypass duct defined between said fan case and a splitter wall of said inner core housing, and a rigid connection between said fan case and said inner core housing including a plurality of A-frames each including a pair of legs rigidly connected at a connection point to said fan case, and each leg in said pair extending away from said connection point in opposed circumferential directions to be connected to a compressor wall of said inner core housing to form an A-shape;

the rigid connection also including a plurality of fan exit guide vanes rigidly connected to said fan case; and

wherein a moment stiffness is defined about an axis extending perpendicular to a rotational axis of the gas turbine engine, and a moment stiffness ratio of a moment stiffness of the plurality of fan exit guide vanes to a moment stiffness of the combination of the plurality of A-frames, the compressor wall, and a fan intermediate case forward of the lower pressure compressor being greater than or equal to 0.50 and less than or equal to 2.0.

18. The gas turbine engine as set forth in claim 17 , wherein the moment stiffness ratio being less than or equal to 1.4.

19. The gas turbine engine as set forth in claim 18 , wherein the moment ratio being less than or equal to 1.0.

20. The gas turbine engine as set forth in claim 17 , wherein a ratio of a quantity of the plurality of fan exit guide vanes to a quantity of the legs of the plurality of A-frames being greater than or equal to 1.75.