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US 8,726,635 B1Granted
Gas turbine engine with dual compression rotor
AssigneeThe United States of America as represented by the Secretary of the Air Force
InventorMark R. Dale
Life of this patent
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Abstract
The present invention provides a gas turbine engine having a combustion chamber section substantially forward of an axial compressor section. An example embodiment uses a centrifugal compressor section behind the axial compressor section to help route compressed air exiting the axial compressor section forward to the combustion chamber section.
Claims1
A gas turbine engine having a front intake section and a rear exhaust section, comprising:
- (a) an axial compressor section between the front intake section and the rear exhaust section;
- (b) a combustion chamber section operatively interconnected with and substantially forward of the axial compressor section, wherein the axial compressor section includes a dual compression rotor, the dual compression rotor comprising two concentric rings of compression blades on a single rotor;
- and
- (c) a centrifugal compressor section behind the axial compressor section.
1 more element
Dependencies are inferred from the claim text.