Gas turbine engine vanes

1. A gas turbine engine defining axial, radial and circumferential directions and comprising:

a duct defined by an axially extending radially inner wall and an axially extending radially outer wall, wherein:

an annular array of turning vanes is provided in the duct, the array comprising a set of cambered aerodynamic vanes and a set of cambered strut-vanes, each of the turning vanes having leading edges and trailing edges;

the leading edges of the strut-vanes are axially upstream of the trailing edges of the aerodynamic vanes;

the trailing edges of the strut-vanes are axially downstream of the trailing edges of the aerodynamic vanes;

a tangent (θ 3 ) to a mean camber line at the trailing edges of the strut-vanes is more closely aligned to the axial direction than is a tangent (θ 2 ) to a mean camber line at the trailing edges of the aerodynamic vanes; and

the mean camber line of each of the strut-vanes is the same as the mean camber lines of at least some of the aerodynamic vanes for axial locations between the leading and trailing edges of the aerodynamic vanes;

wherein, progressing towards the respective trailing edge, the mean camber line of each of the strut-vanes turns towards the axial direction by an angle in the range of 5 degrees and 30 degrees between the axial position of the trailing edges of the aerodynamic vanes and the trailing edge of each of the strut-vanes.

2. The gas turbine engine according to claim 1 , further comprising:

a first compressor section; and

a second compressor section that is axially downstream of the first compressor section, wherein:

the duct is provided between the first compressor section and the second compressor section, with each of the radially inner wall and radially outer wall extending axially downstream and radially inboard as the duct extends away from the first compressor section towards the second compressor section.

3. The gas turbine engine according to claim 2 , wherein:

the first compressor section comprises rotor blades that are connected to a first shaft that is also connected to a first turbine;

the second compressor section comprises rotor blades that are connected to a second shaft that is also connected to a second turbine; and

in use, the rotational speed of the second shaft is higher than the rotational speed of the second shaft.

4. The gas turbine engine according to claim 2 , wherein:

the first compressor section comprises at least two stages, each stage of the first compressor section comprising a rotor row and a stator row; and

the second compressor section comprises at least two stages, each stage of the second compressor section comprising a rotor row and a stator row, wherein:

immediately upstream of the annular array of turning vanes is a rotor row of one of the stages of the first compressor section; and

immediately downstream of the annular array of turning vanes is a rotor row of one of the stages of the second compressor section.

5. The gas turbine engine according to claim 1 , wherein the leading edges of the aerodynamic vanes have an axial position that is aligned with an axial position of the leading edge of the strut-vanes.

6. The gas turbine engine according to claim 1 , wherein the tangent (θ 2 ) to the mean camber line of each of the aerodynamic vanes at the trailing edge is in the range of 5 degrees to 40 degrees from the axial direction.

7. The gas turbine engine according to claim 1 , wherein the tangent to the mean camber line (θ 3 ) of each of the strut-vanes at the trailing edge is in the range of 0 degrees to 15 degrees from the axial direction.

8. The gas turbine engine according to claim 1 , wherein only the set of cambered aerodynamic vanes and the set of cambered strut-vanes extend between the radially inner wall and radially outer wall of the duct.

9. The gas turbine engine according to claim 1 , wherein:

a plurality of the aerodynamic vanes are provided in a gap defined by a suction surface of one of the strut-vanes and a pressure surface of another one of the strut-vanes, the plurality of the aerodynamic vanes comprising at least two aerodynamic vanes that are closest to said suction surface and at least two aerodynamic vanes that are closest to said pressure surface; and

a solidity of the at least two aerodynamic vanes in the gap that are closest to the suction surface is higher than a solidity of the at least two aerodynamic vanes in the gap that are closest to the pressure surface.

10. The gas turbine engine according to claim 9 , wherein a spacing (p) between the two aerodynamic vanes that are closest to said suction surface is greater than a spacing between the two aerodynamic vanes that are closest to said pressure surface.

11. The gas turbine engine according to claim 9 , wherein a chord length (q) of an aerodynamic vane that is closest to said suction surface is smaller than a chord length of an aerodynamic vane that is closest to said pressure surface.

12. The gas turbine engine according to claim 9 , wherein:

the tangent to the mean camber line at the trailing edge of an aerodynamic vane of the plurality of aerodynamic vanes that is closest to said pressure surface is closer to the axial direction than is the tangent to the mean camber line at the trailing edge of an aerodynamic vane of the plurality of aerodynamic vanes that is closest to said suction surface.

13. The gas turbine engine according to claim 1 , wherein at least one of the strut-vanes has components of an auxiliary gas turbine engine system passing therethrough.

14. The gas turbine engine according to claim 1 , wherein the strut-vanes have a maximum thickness (t) in the range of 10% to 30% of their chord length.

15. The gas turbine engine according to claim 1 , wherein a chordwise location of a maximum thickness position of each of the strut-vanes is in the range of 20% to 50% of a chord length thereof.

16. The gas turbine engine according to claim 1 , wherein a thickness of the trailing edges of both the aerodynamic vanes and the strut-vanes is less than 5 mm.

17. A gas turbine engine defining axial, radial and circumferential directions and comprising:

a duct defined by an axially extending radially inner wall and an axially extending radially outer wall, wherein:

an annular array of turning vanes is provided in the duct, the array comprising a set of cambered aerodynamic vanes and a set of cambered strut-vanes, each of the turning vanes having leading edges and trailing edges;

the leading edges of the strut-vanes are axially upstream of the trailing edges of the aerodynamic vanes;

the trailing edges of the strut-vanes are axially downstream of the trailing edges of the aerodynamic vanes;

a tangent (θ 3 ) to a mean camber line at the trailing edges of the strut-vanes is more closely aligned to the axial direction than is a tangent (θ 2 ) to a mean camber line at the trailing edges of the aerodynamic vanes; and

the mean camber line of each of the strut-vanes is the same as the mean camber lines of at least some of the aerodynamic vanes for axial locations between the leading and trailing edges of the aerodynamic vanes;

wherein a plurality of the aerodynamic vanes are provided in a gap defined by a suction surface of one of the strut-vanes and a pressure surface of another one of the strut-vanes, the plurality of the aerodynamic vanes comprising at least one aerodynamic vane that is closest to said suction surface and at least one aerodynamic vane that is closest to said pressure surface;

wherein a chord length (q) of the aerodynamic vane that is closest to said suction surface is smaller than a chord length of the aerodynamic vane that is closest to said pressure surface.