Radial strut frame connection between fan case and core housing in a gas turbine engine

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 a core engine with a compressor section, including a low pressure compressor;

said fan rotor delivering air into a bypass duct defined between said fan case and said core engine, and a rigid connection between said fan case and said core engine including a plurality of radial struts rigidly connected to said fan case, and to said core engine;

a plurality of non-structural fan exit guide vanes and said non-structural fan exit guide vanes being provided with an acoustic feature to reduce noise, said non-structural fan exit guide vanes being rigidly mounted to at least one of the fan case and the core engine;

wherein there are also structural fan exit guide vanes rigidly secured to both said fan case and said core engine; and

wherein said structural fan exit guide vanes extend at a first angle from a radially inner end to a radially outer end, and said first angle having a radially outward component and an axially aft component, and said radial struts extending from a radially inner connection to said core engine to a connection point with said fan case at a second angle having a component in a radially outward direction, and in an axially forward direction.

2. The gas turbine engine as set forth in claim 1 , wherein said acoustic feature includes said non-structural fan exit guide vanes being formed with chambers and a covering perforated face sheet.

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

4. The gas turbine engine as set forth in claim 3 , wherein said core engine includes a fan intermediate case forward of said low pressure compressor, and said structural fan guide vanes are rigidly connected to said fan intermediate case.

5. The gas turbine engine as set forth in claim 4 , wherein said core engine including a compressor intermediate case intermediate said low pressure compressor and a high pressure compressor, and said radial struts rigidly secured to said compressor intermediate case.

6. The gas turbine engine as set forth in claim 4 , wherein said non-structural fan exit guide vanes also being rigidly mounted to said fan intermediate case, but being allowed to float in a radial direction relative to the fan case.

7. The gas turbine engine as set forth in claim 3 , wherein said acoustic feature includes said non-structural fan exit guide vanes being formed with chambers and a covering perforated face sheet.

8. The gas turbine engine as set forth in claim 1 , wherein said core engine includes a fan intermediate case forward of said low pressure compressor, and said structural fan guide vanes are rigidly connected to said fan intermediate case.

9. The gas turbine engine as set forth in claim 8 , wherein said structural fan exit guide vanes include 15-30% of the total fan exit guide vanes.

10. The gas turbine engine as set forth in claim 8 , wherein said low pressure compressor has four to six stages.

11. The gas turbine engine as set forth in claim 1 , wherein said radial struts are circumferentially spaced by a uniform spacing distance about a circumference of the fan case and core engine.

12. The gas turbine engine as set forth in claim 11 , wherein there is one of said radial struts at a top dead center position and one of said radial struts at a bottom dead center position.

13. The gas turbine engine as set forth in claim 12 , wherein there are at least two radial struts circumferentially intermediate the radial strut at top dead center and said radial strut at bottom dead center.

14. The gas turbine engine as set forth in claim 13 , wherein said radial struts at top and bottom dead center are hollow bifurcations.

15. The gas turbine engine as set forth in claim 13 , wherein said low pressure compressor has four to six stages.

16. The gas turbine engine as set forth in claim 1 , wherein said structural fan exit guide vanes include greater than 0 and less than 55% of a total number of fan exit guide vanes including the non-structural fan exit guide vanes.

17. The gas turbine engine as set forth in claim 16 , wherein said low pressure compressor has four to six stages.

18. The gas turbine engine as set forth in claim 16 , wherein said structural fan exit guide vanes include 15-30% of the total fan exit guide vanes.