Geared gas turbine engine with front section moment stiffness relationships

Abstract

A gas turbine engine includes a fan rotor driven by a fan drive turbine about an axis through a gear reduction. An inner core engine has an inner core engine housing surrounding a compressor section, including a low pressure compressor. A rigid connection between a fan case and the inner core engine includes A-frames rigidly connected at a connection point to the fan case. Fan exit guide vanes rigidly connect to the fan case, and to the inner core engine. A fan intermediate case is positioned forward of a first rotor stage in the low pressure compressor. A rigid structure is connected to the inner core engine and to the fan exit guide vanes. The rigid structure defines a structure moment stiffness. The fan intermediate case defines an intermediate case moment stiffness. A ratio of the structure moment stiffness to the intermediate case moment stiffness is between 5 and 15.

What is claimed is: 
     
         1 . A gas turbine engine comprising:
 a compressor section having a low pressure compressor, a high pressure compressor and a compressor intermediate case;   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 said compressor section and a turbine section; and   a first mount for said engine attached to one of said fan case or said compressor intermediate case, and a second mount for said engine attached to said inner core housing, said first and second mount for attaching the engine to an aircraft, and an axial length of said engine defined between a forwardmost point of said engine and a rearwardmost point of said engine and a mount distance being defined between a trailing edge of said first mount and a trailing edge of said second mount, with each of said trailing edges measured from a radially innermost end, and a ratio of said mount distance to said axial length being between 0.4 and 0.7 if said first mount is attached to said fan case, and being between 0.4 and 0.6 when said first mount is attached to said compressor intermediate case.   
     
     
         2 . The gas turbine engine as set forth in  claim 1 , wherein said low pressure compressor has four to six stages. 
     
     
         3 . The gas turbine engine as set forth in  claim 2 , wherein there is a high pressure turbine driving a high pressure compressor, and a mid-turbine frame defined between said high pressure turbine and said fan drive turbine, and said second mount being attached to said inner core housing at a location aligned with said mid-turbine frame. 
     
     
         4 . The gas turbine engine as set forth in  claim 3 , wherein a plurality of fan exit guide vanes connect said fan case to said inner core housing, and said first mount being on said fan case at an axial location aligned with a radially outermost edge of said plurality of fan exit guide vanes. 
     
     
         5 . The gas turbine engine as set forth in  claim 3 , wherein said first mount is attached to said compressor intermediate case. 
     
     
         6 . The gas turbine engine as set forth in  claim 3 , wherein there are 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 said inner core engine. 
     
     
         7 . The gas turbine engine as set forth in  claim 1 , wherein there is a high pressure turbine driving a high pressure compressor, and a mid-turbine frame defined between said high pressure turbine and said fan drive turbine, and said second mount being attached to said inner core housing at a location aligned with said mid-turbine frame. 
     
     
         8 . The gas turbine engine as set forth in  claim 7 , wherein a plurality of fan exit guide vanes connect said fan case to said inner core housing, and said first mount being on said fan case at an axial location aligned with a radially outermost edge of said plurality of fan exit guide vanes. 
     
     
         9 . The gas turbine engine as set forth in  claim 7 , wherein said first mount is attached to said compressor intermediate case. 
     
     
         10 . The gas turbine engine as set forth in  claim 7 , wherein there are 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 said inner core engine. 
     
     
         11 . A gas turbine engine comprising:
 a compressor section having a low pressure compressor, a high pressure compressor and a compressor intermediate case;   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 said compressor section and a turbine section; and   a first mount for said engine attached to said fan case, and a second mount for said engine attached to said inner core housing, said first and second mount for attaching the engine to an aircraft, and an axial length of said engine defined between a forwardmost point of said engine and a rearwardmost point of said engine and a mount distance being defined between a trailing edge of said first mount and a trailing edge of said second mount, with each of said trailing edges measured from a radially innermost end, and a ratio of said mount distance to said axial length being between 0.4 and 0.7.   
     
     
         12 . The gas turbine engine as set forth in  claim 11 , wherein said low pressure compressor has four to six stages. 
     
     
         13 . The gas turbine engine as set forth in  claim 11 , wherein there is a high pressure turbine driving a high pressure compressor, and a mid-turbine frame defined between said high pressure turbine and said fan drive turbine, and said second mount being attached to said inner core housing at a location aligned with said mid-turbine frame. 
     
     
         14 . The gas turbine engine as set forth in  claim 11 , wherein a plurality of fan exit guide vanes connect said fan case to said inner core housing, and said first mount being on said fan case at an axial location aligned with a radially outermost edge of said plurality of fan exit guide vanes. 
     
     
         15 . The gas turbine engine as set forth in  claim 11 , wherein there are 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 said inner core engine. 
     
     
         16 . A gas turbine engine comprising:
 a compressor section having a low pressure compressor, a high pressure compressor and a compressor intermediate case;   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 said compressor section and a turbine section; and   a first mount for said engine attached to said compressor intermediate case, and a second mount for said engine attached to said inner core housing, said first and second mount for attaching the engine to an aircraft, and an axial length of said engine defined between a forwardmost point of said engine and a rearwardmost point of said engine and a mount distance being defined between a trailing edge of said first mount and a trailing edge of said second mount, with each of said trailing edges measured from a radially innermost end, and a ratio of said mount distance to said axial length being between 0.4 and 0.6.   
     
     
         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 17 , wherein there is a high pressure turbine driving a high pressure compressor, and a mid-turbine frame defined between said high pressure turbine and said fan drive turbine, and said second mount being attached to said inner core housing at a location aligned with said mid-turbine frame. 
     
     
         19 . The gas turbine engine as set forth in  claim 16 , wherein there is a high pressure turbine driving a high pressure compressor, and a mid-turbine frame defined between said high pressure turbine and said fan drive turbine, and said second mount being attached to said inner core housing at a location aligned with said mid-turbine frame. 
     
     
         20 . The gas turbine engine as set forth in  claim 16 , wherein there are 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 said inner core engine.