Radial cooling system for gas turbine engine compressors

A compressor for a gas turbine engine, comprising:

• a first rotor supporting a first plurality of circumferentially spaced rotor blades, a second rotor disposed axially downstream of the first rotor and supporting a second plurality of circumferentially spaced rotor blades and a third rotor disposed axially downstream of the second rotor and supporting a third plurality of circumferentially spaced rotor blades;
• a first bore cavity between the first rotor and the second rotor, a second bore cavity between the second rotor and the third rotor and a third bore cavity between the third rotor and an aft hub;
• a fluid passageway configured to provide cooled air to the first bore cavity, the second bore cavity and the third bore cavity;
• a first anti-vortex component positioned within the first bore cavity and configured to increase pressure of the cooled air as the cooled air traverses radially outward from the first bore cavity;
• a second anti-vortex component positioned within the second bore cavity and configured to increase pressure of the cooled air as the cooled air traverses radially outward from the second bore cavity;
• and
• a third anti-vortex component positioned within the third bore cavity and configured to increase pressure of the cooled air as the cooled air traverses radially outward from the third bore cavity;
• and
• an orifice through the aft hub configured to introduce an aft cooled air flow path from the third bore cavity onto a downstream rim portion of the third rotor,
• wherein the downstream rim portion of the third rotor is disposed axially upstream and adjacent a plurality of exit guide vanes.