Gas turbine compression system and compressor structure

• US 8,757,965 B2
• Filed: 06/02/2005
• Issued: 06/24/2014
• Est. Priority Date: 06/01/2004
• Status: Active Grant

• Current Assignee: GKN Aerospace Sweden AB (Melrose Industries Plc)
• Original Assignee: Volvo Aero Corporation (Melrose Industries Plc)
• Inventors: Baralon, Stephane
• Primary Examiner(s): VERDIER, CHRISTOPHER M
• Application Number:
  US10/908,970
• Publication Number:
  US 20060275110A1
• Time in Patent Office: 3,309 Days
• Field of Search: 415/191, 415/193, 415/194, 415/195, 415/142, 415/208.2, 415/209.1, 415/211.2, 415/199.5, 602/261
• US Class Current: 415/142
• CPC Class Codes: F01D 25/162 Bearing supports
• CPC Class Codes: F01D 9/041 using blades F01D5/148 takes precedence
• CPC Class Codes: F01D 9/065 Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids see also F01D25/16, F01D25/24 and F01D25/26
• CPC Class Codes: F05D 2260/961 by mistuning rotor blades or stator vanes with irregular interblade spacing, airfoil shape
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A gas turbine compression system (1) comprising a gas channel (5), a low pressure compressor section (8) and a high pressure compressor section (9) for compression of the gas in the channel and a compressor structure (14) arranged between the low pressure compressor section (8) and the high pressure compressor section (9). The compressor structure (14) is designed to conduct a gas flow in the gas channel and comprises a plurality of radial struts (15,16,21,24,25) for transmission of load, wherein at least one of the struts (15,16,21,24,25) is hollow for housing service components. The compressor structure (14) is arranged directly downstream of a last rotor (10) in the low pressure compressor section (8) and designed for substantially turning a swirling gas flow from the rotor (10) by a plurality of the struts (15,16,21,24,25) having a cambered shape.

Claims

1. A gas turbine compression system (1), comprising:

a gas channel (5),

a low pressure compressor section (8) and a high pressure compressor section (9) for compression of gas in the channel, the low pressure compressor section (8) comprising an alternating series of rotating rotor blades and stationary stator blades with the last blades in the series being rotor blades, and

a compressor structure (14) arranged between the low pressure compressor section (8) and the high pressure compressor section (9),

wherein the compressor structure (14) is designed to conduct a gas flow in the gas channel and comprises a plurality of radial struts (15, 16, 21, 24, 25) for transmission of load, with the struts being spaced by a distance downstream from the last low-pressure-compressor rotor blades in the series with no intervening stator blades therebetween,

wherein at least one of said struts is hollow for housing service components, and

wherein the compressor structure (14) is designed for substantially turning a swirling gas flow from the last low-pressure-compressor rotor blades in the series by virtue of a plurality of said struts (15, 16, 21, 24, 25) having a cambered shape.

2. The gas turbine compression system according to claim 1, wherein the direction of a mean camber line (M) at the leading edge (101) of at least one of the cambered struts (15, 16, 21, 24, 25) is inclined by at least 20°

in relation to the direction of the mean camber line (M) at the trailing edge (102) of said cambered strut.

3. The gas turbine compression system according to claim 1, wherein the direction of a mean camber line (M) at the leading edge (101) of at least one of the cambered struts (15, 16, 21, 24, 25) is inclined by at least 30°

in relation to the direction of the mean camber line (M) at the trailing edge (102) of said cambered strut.

4. The gas turbine compression system according to claim 1, wherein the direction of a mean camber line (M) at the leading edge (101) of at least one of the cambered struts (15, 16, 21, 24, 25) is inclined by at least 40°

in relation to the direction of the mean camber line (M) at the trailing edge (102) of said cambered strut.

5. The gas turbine compression system according to claim 1, wherein the thickness to chord ratio of at least one of the cambered struts (15, 16, 21, 24, 25) is about 0.10.

6. The gas turbine compression system according to claim 1, wherein the struts (15, 16, 21, 24, 25) are arranged so that the leading edge (101) of each of them is located at substantially the same position in a direction parallel to the rotational axis (18) of the gas turbine compression system.

7. The gas turbine compression system according to claim 1, wherein the compressor structure (14) comprises a plurality of aerodynamic vanes (19, 20) with a substantially smaller cross section area relative to the struts (15, 16, 21, 24, 25).

8. The gas turbine compression system according to claim 7, wherein at least one of said aerodynamic vanes (19, 20) is located between two adjacent struts (16,21) in the circumferential direction of the compressor structure (14).

9. The gas turbine compression system according to claim 7, wherein a leading edge of each of a first set of said aerodynamic vanes (19, 20) is arranged at substantially the same position as the leading edge of the struts (15, 16, 21, 24, 25) in a direction parallel to the rotational axis (18) of the gas turbine compression system.

10. The gas turbine compression system according to claim 7, wherein at least one of said aerodynamic vanes (19, 20) is substantially solid in cross section.

11. The gas turbine compression system according to claim 1, wherein said struts (15, 16, 21, 24, 25) are asymmetrically distributed in the circumferential direction of the compressor structure.

12. The gas turbine compression system according to claim 1, wherein said struts (15, 16, 21, 24, 25) are symmetrically distributed with regard to a plane in parallel with the axial direction of the gas turbine compression system.

13. The gas turbine compression system according to claim 1, wherein the compressor structure (14) is arranged directly upstream a first rotor (11) in the high pressure compressor section.