Abstract: A gas turbine engine for an aircraft includes a compressor section where at least one of the airfoil members defines a vane exit vector extending tangentially from a curved surface of the airfoil member adjacent a trailing edge of the airfoil member, a projection of the vane exit vector in a longitudinal plane perpendicular to a radial direction of the engine extending at an airfoil angle from the longitudinal axis. A bleed slot defined through the casing wall and providing fluid communication between the core air passage and the bleed duct extends along a slot axis. A projection of the slot axis in the longitudinal plane extends at a slot angle with respect to the longitudinal axis. The slot angle is different from the airfoil angle. A method of forming bleed slots in a gas turbine engine is also discussed.