Compressor bleed pressure recovery

A gas turbine engine comprising:

• a compressor section including (a) a compressor rotor shaft assembly, (b) a stator shroud casing surrounding the compressor rotor shaft assembly, a compressor flow passage being defined between the compressor rotor shaft assembly and the stator shroud casing, and (c) a plurality of bleed section stator vanes extending from the stator shroud casing into the compressor flow passage, a leading edge plane being defined perpendicular to a compressor centerline axis at a leading edge of at least one of the plurality of bleed section stator vanes;
• and
• a compressor bleed air system including (a) a compressor bleed air duct including an inlet end and an outlet end, the inlet end extending through the stator shroud casing and providing airflow communication between the compressor flow passage and the compressor bleed air duct, and (b) a plurality of submerged sloped inlet portions in the stator shroud casing each extending upstream from the inlet end of the compressor bleed air duct to at least the leading edge plane, each of the plurality of submerged sloped inlet portions expanding radially outward in the stator shroud casing from an upstream end of the submerged sloped inlet portion to the inlet end of the compressor bleed air duct, the inlet end of the compressor bleed air duct being arranged downstream of the plurality of bleed section stator vanes,
• wherein the plurality of bleed section stator vanes are circumferentially spaced about the stator shroud casing, and respective ones of the plurality of submerged sloped inlet portions are circumferentially arranged between a respective pair of the plurality of bleed section stator vanes.