Gas turbine engine and method for cooling the compressor of a gas turbine engine

• US 8,979,470 B2
• Filed: 08/04/2011
• Issued: 03/17/2015
• Est. Priority Date: 08/10/2010
• Status: Active Grant

• Current Assignee: Ansaldo Energia Switzerland AG
• Original Assignee: Alstom Technology Ltd. (Alstom SA)
• Inventors: Riazantsev, Sergei, Kiewel, Holger, Olmes, Sven, Kramer, Thomas, Shchukin, Sergey
• Primary Examiner(s): Kershteyn, Igor
• Application Number:
  US13/197,840
• Publication Number:
  US 20120036864A1
• Time in Patent Office: 1,321 Days
• Field of Search: 415/1, 415/115, 415/116, 415/175, 415/176
• US Class Current: 415/1
• CPC Class Codes: F01D 5/084 the fluid circulating at the periphery of a multistage rotor, e.g. of drum type
• CPC Class Codes: F01D 5/3007 of axial insertion type
• CPC Class Codes: F04D 29/321 for axial flow compressors
• CPC Class Codes: F04D 29/584 cooling or heating the machine F04D29/5846, F04D29/5853 take precedence
• CPC Class Codes: F05D 2220/32 in gas turbines

Abstract

A gas turbine engine includes a compressor with rotor blades having roots connected into seats of a compressor drum. The rotor blade roots and/or the compressor drum have longitudinal passages for a cooling fluid, connecting higher pressure areas to lower pressure areas of the gas turbine engine.

Claims

1. A gas turbine engine, comprising:

a compressor including a compressor drum and rotor blades having roots connected into seats of the compressor drum, wherein at least one of the rotor blade roots and the compressor drum include longitudinal passages for a cooling fluid, the longitudinal passages connecting higher pressure areas to lower pressure areas of the gas turbine engine;

a gap downstream of the compressor drum for separating the compressor drum from a combustion chamber; and

a protrusion provided within the gap to close a compressed air path, wherein the higher pressure areas are defined between the protrusion and the compressed air path.

2. The gas turbine engine as claimed in claim 1, wherein the seats are defined by longitudinal slots into which the blade roots are inserted.

3. The gas turbine engine as claimed in claim 2, wherein the rotor blade roots include the longitudinal passages defined by longitudinal channels provided in the blade roots, wherein channels of blade roots inserted into the same seat are connected together.

4. The gas turbine engine as claimed in claim 3, comprising:

spacers between two adjacent blade roots inserted into the same seat, the spacers having a spacer root and a platform defining, with platforms of the rotor blades, a compressed air path, wherein the spacer roots have longitudinal passages connected to the passages of the blade roots.

5. The gas turbine engine as claimed in claim 1, wherein the lower pressure areas are defined by areas of the gap below the protrusion.

6. The gas turbine engine as claimed in claim 1, wherein the compressor drum is hollow, and the lower pressure areas are defined in the inside of the hollow compressor drum.

7. The gas turbine engine as claimed in claim 1, comprising:

a circumferential chamber extending at an intermediate position of the compressor drum, the circumferential chamber being connected to at least one of the longitudinal passages of the blade roots and to the longitudinal passages of the compressor drum.

8. The gas turbine engine as claimed in claim 3, comprising:

each of the blade roots and compressor drum having longitudinal passages, wherein the longitudinal passages of the blade roots and the longitudinal passages of the rotor drum have axes parallel to an engine longitudinal axis and have a same radial distance from it.

9. The gas turbine engine as claimed in claim 8, wherein the longitudinal passages of the blade roots are connected to the lower pressure areas and the longitudinal passages of the compressor drum are connected to the higher pressure areas.

10. A method for cooling a compressor of a gas turbine engine, including a compressor with rotor blades having roots connected into seats of a compressor drum, the method comprising:

forming at least one of the blade roots and the compressor drum with longitudinal passages for a cooling fluid, the longitudinal passages connecting higher pressure areas to lower pressure areas of the gas turbine engine;

passing a cooling fluid through the longitudinal passages;

forming a gap downstream of the compressor drum for separating the compressor drum from a combustion chamber; and

providing a protrusion within the gap for closing a compressed air path, for defining the higher pressure areas between the protrusion and the compressed air path.

11. A gas turbine engine, comprising:

a compressor including a compressor drum and rotor blades having roots connected into seats of the compressor drum, wherein the compressor drum and at least one rotor blade root each include longitudinal passages for a cooling fluid, the longitudinal passages connecting higher pressure areas to lower pressure areas of the gas turbine engine,

wherein at least one of the longitudinal passages in the compressor and at least one of the longitudinal passages in the rotor blade roots have an axis parallel to an engine longitudinal axis and have the same radial distance from the axis.

12. The gas turbine engine of claim 11, wherein a compressor drum longitudinal passage is connected to a rotor blade root longitudinal passage, said longitudinal passages being parallel and adjacent in a circumferential direction to one another.

13. The gas turbine engine of claim 11, wherein at least one of the longitudinal passages in the compressor and at least one of the longitudinal passages in the rotor blade roots are connected by a circumferential chamber.