A gas turbine engine comprising: a fan including a plurality of fan blades rotatable about an engine axis (A), 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 (E) that is substantially transverse to the engine axis (A); 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 (A) 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 (E) 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 (CD) 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 (A) 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), 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 (E) that is substantially transverse to the engine axis (A); 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 (A) 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 (E) 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 (CD) 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 (A) 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. <EOS>