Dual compression rotor

• US 8,192,141 B1
• Filed: 04/07/2008
• Issued: 06/05/2012
• Est. Priority Date: 04/05/2007
• Status: Expired due to Fees

• Current Assignee: United States Of America As Represented By The Secretary Of The Air Force
• Original Assignee: United States Of America As Represented By The Secretary Of The Air Force
• Inventors: Dale, Mark R.
• Primary Examiner(s): EDGAR, RICHARD A
• Application Number:
  US12/156,808
• Time in Patent Office: 1,520 Days
• Field of Search: 415/198.1, 415/199.4, 415/199.5, 415/79, 416/175, 416/203, 416/199
• US Class Current: 415/79
• CPC Class Codes: F02C 3/073 the compressor and turbine stages being concentric
• CPC Class Codes: F02C 3/145 the combustion chamber being in the reverse flow-type
• CPC Class Codes: F02K 3/068 being characterised by a short axial length relative to the diameter
• CPC Class Codes: F04D 1/08 the stages being situated concentrically
• CPC Class Codes: F04D 19/022 with concentric rows of vanes;
• CPC Class Codes: F04D 25/045 the pump wheel carrying the fluid driving means, e.g. turbine blades
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

The present invention provides a gas turbine rotor having a first axial compression stage for compressing air in a first direction and a second axial compression stage for compressing the air in a second direction, the second direction being generally opposite the first direction.

Claims

1. A gas turbine rotor comprising:

a plurality of first axial compression blades configured for compressing air in a first direction, the plurality of first axial compression blades having a first blade quantity and a first blade height; and

a plurality of second axial compression blades configured for compressing the air from the first axial compression blades in a second direction, the second direction being generally opposite the first direction, the plurality of second axial compression blades having a second blade quantity that is greater than the first blade quantity and a second blade height that is less than the first blade height, wherein the plurality of second axial compression blades are discontinuous with the plurality of first axial compression blades and extend radially outward from the plurality of first axial compression blades such that the plurality of second axial compression blades are substantially coplanar with the plurality of first axial compression blades.

2. The rotor of claim 1 further comprising an intermediate ring connected to outer radial ends of the plurality of first axial compression blades and connected to inner radial ends of the plurality of second axial compression blades.

3. The rotor of claim 2 wherein the intermediate ring is a barrier between the air compressed by the plurality of first axial compression blades and the air compressed by the plurality of second axial compression blades.

4. The rotor of claim 3 further comprising a rim connected to outer radial ends of the plurality of second axial compression blades.

5. The rotor of claim 4 further comprising a plurality of turbine blades, wherein the plurality of turbine blades are discontinuous with the plurality of second axial compression blades such that the plurality of turbine blades have a third blade quantity that is greater than the second blade quantity and a third blade height that is greater than the second blade height, the inner radial ends of the plurality of turbine blades being connected to the rim.

6. The rotor of claim 5 wherein the plurality of turbine blades is configured such that the rotor rotates clockwise as exhaust gas passes over the plurality of turbine blades.

7. The rotor of claim 5 wherein the plurality of turbine blades is configured such that the rotor rotates counterclockwise as exhaust gas passes over the plurality of turbine blades.

8. The rotor of claim 5 wherein the rim is a barrier between the air compressed by the plurality of second axial compression blades and exhaust gas traveling over the plurality of turbine blades.

9. A gas turbine rotor comprising:

an inner ring;

a plurality of first axial compression blades connected to the inner ring, the plurality of first axial compression blades having a first blade quantity and a first blade height;

an intermediate ring connected to the plurality of first axial compression blades;

a plurality of second axial compression blades connected to the intermediate ring, the plurality of second axial compression blades having a second blade quantity that is greater than the first blade quantity and a second blade height that is less than the first blade height, wherein the plurality of second axial compression blades are discontinuous with the plurality of first axial compression blades;

a rim connected to the plurality of second axial compression blades; and

a plurality of turbine blades connected to the rim, the plurality of turbine blades having a third blade quantity that is greater than the second blade quantity and a third blade height that is greater than the second blade height, wherein the plurality of turbine blades are discontinuous with the plurality of second axial compression blades.

10. The rotor of claim 9 wherein the plurality of first axial compression blades compress air axially in a backward direction when the rotor rotates.

11. The rotor of claim 10 wherein the plurality of second axial compression blades compress the air in a forward direction when the rotor rotates.

12. The rotor of claim 11 wherein the plurality of turbine blades is configured for rotating the rotor when exhaust gas passes by the plurality of turbine blades.