A gas turbine engine (20;100;120;140;150;170) comprising: a fan rotor (42;174) having fan blades (98;134) received within an outer nacelle (94;122;145;172), and said outer nacelle (94... 172) having an inner surface (132;151); a distance (L) being defined between an axial outer end (B) of said nacelle (94... 172) and a leading edge of said fan blades (98;134); and an anti-icing treatment provided to an inner periphery of said nacelle (94... 172) over at least 75% of the distance along said inner periphery of said nacelle (94... 172), wherein the distance (L) is defined from a plane (X) defined by leading edges of said fan blades (98;134) to an axial location of a forwardmost part of said nacelle (94... 172), and an outer diameter (D) of said fan blades (98;134) is defined; characterised in that : a ratio of said distance (L) to said outer diameter (D) is between 0.2 and 0.5; said outer end (B) of said nacelle (94... 172) extends outwardly for varying extents across a circumference of said nacelle (94... 172), and said ratio of said distance (L) to the outer diameter (D) for all locations of said nacelle (94... 172) is between 0.2 and 0.45; a fan drive turbine drives said fan rotor (42;174) through a gear reduction ; and said nacelle (94... 172) 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 location centered substantially about bottom dead center when the gas turbine engine (20... 170) is mounted on an aircraft, and an anti-icing feature being mounted where the acoustic treatment is not provided. a fan rotor (42;174) having fan blades (98;134) received within an outer nacelle (94;122;145;172), and said outer nacelle (94... 172) having an inner surface (132;151); a distance (L) being defined between an axial outer end (B) of said nacelle (94... 172) and a leading edge of said fan blades (98;134); and an anti-icing treatment provided to an inner periphery of said nacelle (94... 172) over at least 75% of the distance along said inner periphery of said nacelle (94... 172), wherein the distance (L) is defined from a plane (X) defined by leading edges of said fan blades (98;134) to an axial location of a forwardmost part of said nacelle (94... 172), and an outer diameter (D) of said fan blades (98;134) is defined; a ratio of said distance (L) to said outer diameter (D) is between 0.2 and 0.5; said outer end (B) of said nacelle (94... 172) extends outwardly for varying extents across a circumference of said nacelle (94... 172), and said ratio of said distance (L) to the outer diameter (D) for all locations of said nacelle (94... 172) is between 0.2 and 0.45; a fan drive turbine drives said fan rotor (42;174) through a gear reduction ; and said nacelle (94... 172) 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 location centered substantially about bottom dead center when the gas turbine engine (20... 170) is mounted on an aircraft, and an anti-icing feature being mounted where the acoustic treatment is not provided. <EOS>