A gas turbine engine comprising: a fan including a plurality of fan blades rotatable about an engine axis, a diameter of the fan having a dimension (D 1) that is based on a dimension of the fan blades established between circumferentially outermost edges or tips of the fan blades corresponding to leading edges, each of the fan blades having a leading edge and being rotatable about a fan blade axis that is substantially transverse to the engine axis; and a nacelle assembly arranged at least partially about the fan, the nacelle assembly including an inlet portion forward of the fan and a bypass flow path, a length of the inlet portion having a dimension (L 1) at the engine axis between a location of a plane including a forwardmost part of the leading edge of at least some of the fan blades and a plane including the forwardmost edges on the inlet portion, wherein a dimensional relationship of L 1 /D 1 is less than or equal to 0.4, characterised in that rotation of each of the fan blades about the corresponding fan blade axis is bounded between a first position (P 1) and a second position (P 2) to define a pitch change angle (α) defined relative to a chord extending between the leading edge and a trailing edge of each of the fan blades for the first and second position (P 1, P 2), wherein the first position (P 1) and the second position (P 2) are between a first reference plane (REF 1) extending in a radial direction through the engine axis and a second reference plane (REF 2) perpendicular to the first reference plane (REF 1), wherein the pitch change angle (α) is less than or equal to 60 degrees and is also greater than 20 degrees, wherein a pitch angle change mechanism prohibits the leading edge of the fan blade from rotating through the first reference plane (REF 1) to produce reverse thrust. a fan including a plurality of fan blades rotatable about an engine axis, a diameter of the fan having a dimension (D 1) that is based on a dimension of the fan blades established between circumferentially outermost edges or tips of the fan blades corresponding to leading edges, each of the fan blades having a leading edge and being rotatable about a fan blade axis that is substantially transverse to the engine axis; and a nacelle assembly arranged at least partially about the fan, the nacelle assembly including an inlet portion forward of the fan and a bypass flow path, a length of the inlet portion having a dimension (L 1) at the engine axis between a location of a plane including a forwardmost part of the leading edge of at least some of the fan blades and a plane including the forwardmost edges on the inlet portion, wherein a dimensional relationship of L 1 /D 1 is less than or equal to 0.4, characterised in that rotation of each of the fan blades about the corresponding fan blade axis is bounded between a first position (P 1) and a second position (P 2) to define a pitch change angle (α) defined relative to a chord extending between the leading edge and a trailing edge of each of the fan blades for the first and second position (P 1, P 2), wherein the first position (P 1) and the second position (P 2) are between a first reference plane (REF 1) extending in a radial direction through the engine axis and a second reference plane (REF 2) perpendicular to the first reference plane (REF 1), wherein the pitch change angle (α) is less than or equal to 60 degrees and is also greater than 20 degrees, wherein a pitch angle change mechanism prohibits the leading edge of the fan blade from rotating through the first reference plane (REF 1) to produce reverse thrust.