Gas turbine engine with short inlet, acoustic treatment and anti-icing features

1. A gas turbine engine comprising:

a fan rotor having fan blades received within an outer nacelle, and said outer nacelle having an inner surface; and

a first distance being defined between an axial outer end of said nacelle, and a leading edge of said fan blade, and there being an anti-icing treatment provided to an inner periphery of said nacelle over at least 75% of the first distance along said inner periphery of said nacelle;

wherein a second distance is defined from a plane defined by leading edges of said fan blades to an axial location of a forward most part of said nacelle, and an outer diameter of said fan blades being defined, and a ratio of said second distance to said outer diameter is between 0.2 and 0.5;

wherein said outer end of said nacelle extends outwardly for varying extents across a circumference of said nacelle, and said ratio of said second distance to the outer diameter for all locations of said nacelle being between 0.2 and 0.45;

wherein a fan drive turbine driving said fan rotor through a gear reduction;

wherein said nacelle is provided with an acoustic treatment over the majority of its circumference at an inner periphery, but there being no acoustic treatment across at least 20° at a no acoustic treatment location centered substantially bottom dead center when the gas turbine engine is mounted on an aircraft, and an anti-icing feature being mounted where the acoustic treatment is not provided;

said no acoustic treatment location being at a common axial cross-section as said acoustic treatment, such that said acoustic treatment has circumferential ends and said anti-icing feature extends over only a limited portion of said circumference and is circumferentially intermediate those circumferential ends; and

wherein said anti-icing feature below provided over at least 90% of said no acoustic treatment location.

2. The gas turbine engine as set forth in claim 1 , wherein a tube is positioned within a bulkhead at the outer end of said nacelle, said tube delivering heated air against said outer end of said nacelle, and the air then moving rearwardly to transpiration holes placed adjacent a wall defining an inner end of said bulkhead, with said air being allowed to leave said transpiration holes and move along the inner periphery of said nacelle as the anti-icing treatment.

3. The gas turbine engine as set forth in claim 1 , wherein said anti-icing treatment includes providing a heated circuit along the inner periphery of said nacelle.

4. The gas turbine engine as set forth in claim 3 , wherein said heated circuit includes a foil being provided with a current to generate heat.

5. The gas turbine engine as set forth in claim 1 , wherein said nacelle includes a bulkhead at the outer end and an inner chamber positioned inwardly from said bulkhead, and wherein a tube in said inner chamber delivers air against an inner face of said nacelle as the anti-icing treatment.