An aircraft gas turbine engine comprising: a plurality of rotor blades; a plurality of stator vanes provided rearward of the plurality of rotor blades, characterized in that a stator vane of the plurality of stator vanes comprises: an airfoil body having an airfoil cross section, wherein a maximum airfoil thickness position of the airfoil body in the airfoil cross section satisfies following conditions at least on a tip side of the airfoil body: (a) on a plane expanded in a circumferential direction in which the plurality of stator vanes are arranged, the maximum airfoil thickness position is located in a region between a trailing edge of the airfoil body and an intersection of the airfoil body and a line which is parallel to an extension line of a camber line of a rotor blade of the plurality of rotor blades at a trailing edge of the rotor blade of the plurality of rotor blades and passes through a leading edge of another stator vane of the plurality of stator vanes adjacent in the circumferential direction, and (b) the maximum airfoil thickness position is located in a region between a position of a chord ratio 0.2 and a position of chord ratio 0.8, and the chord ratio is a value obtained by dividing a distance from the leading edge of the airfoil body to a position on a chord of the airfoil body by a full length of the chord. a plurality of rotor blades; a plurality of stator vanes provided rearward of the plurality of rotor blades, an airfoil body having an airfoil cross section, wherein a maximum airfoil thickness position of the airfoil body in the airfoil cross section satisfies following conditions at least on a tip side of the airfoil body: (a) on a plane expanded in a circumferential direction in which the plurality of stator vanes are arranged, the maximum airfoil thickness position is located in a region between a trailing edge of the airfoil body and an intersection of the airfoil body and a line which is parallel to an extension line of a camber line of a rotor blade of the plurality of rotor blades at a trailing edge of the rotor blade of the plurality of rotor blades and passes through a leading edge of another stator vane of the plurality of stator vanes adjacent in the circumferential direction, and (b) the maximum airfoil thickness position is located in a region between a position of a chord ratio 0.2 and a position of chord ratio 0.8, and (a) on a plane expanded in a circumferential direction in which the plurality of stator vanes are arranged, the maximum airfoil thickness position is located in a region between a trailing edge of the airfoil body and an intersection of the airfoil body and a line which is parallel to an extension line of a camber line of a rotor blade of the plurality of rotor blades at a trailing edge of the rotor blade of the plurality of rotor blades and passes through a leading edge of another stator vane of the plurality of stator vanes adjacent in the circumferential direction, and (b) the maximum airfoil thickness position is located in a region between a position of a chord ratio 0.2 and a position of chord ratio 0.8, the chord ratio is a value obtained by dividing a distance from the leading edge of the airfoil body to a position on a chord of the airfoil body by a full length of the chord.