Gas turbine engine

A gas turbine engine for an aircraft comprising:

• an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;
• a fan located upstream of the engine core, the fan comprising a plurality of fan blades;
• and
• a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein:
• for a given cross-section through a fan blade, each fan blade has a camber line defined by the midpoint between its pressure surface and its suction surface, and a covered passage defined as the portion of the cross-section between the trailing edge and a line (J) passing through a point (K) on the suction surface that is closest to the leading edge of a neighbouring fan blade and the leading edge of that neighbouring fan blade;
• at cruise conditions, each cross-section through the blade experiences an inlet relative Mach number M1rel;
• and
• for all cross-sections through each fan blade for which a value of M1rel at cruise is less than 0.9, the change in angle of the camber line between the leading edge (α1) and the point on the camber line that corresponds to the start of the covered passage (α2) satisfies:
• |α2−α1|≥9;
• and/or
• for all cross-sections through each fan blade for which a value of M1rel at cruise is less than 0.75, the change in angle of the camber line between the leading edge (α1) and the point on the camber line that corresponds to the start of the covered passage (α2) satisfies:
• |α2−α1|≥10;
• and/or
• for all cross-sections through each fan blade for which a value of M1rel at cruise is less than 0.8, the change in angle of the camber line between the leading edge (α1) and the point on the camber line that corresponds to the start of the covered passage (α2) satisfies:
• |α2−α1|≥33−(M1rel*30).