Gas turbine compression system and compressor structure
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.
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.