Patent Application
20200325781
The disclosure relates generally to gas turbine engines, and more particularly to an internally cooled turbine blade.
Creep is defined as a time-dependent strain or distortion experienced by materials such as metals when exposed to continued stress and high temperatures. The stress may be in the elastic range below the material yield strength and high temperature may be below the melting point but time-dependent creep strain or deformation results from certain parameters.
Gas turbine blades are exposed to centrifugal stress from high turbine rotational speeds, lateral stresses from gas path flow resistance, and high temperature. Creep distortion of turbine blades can stretch the blade length such that the blade tip interferes with the turbine shroud. Creep can also change the airfoil shape reducing aerodynamic efficiency. Creep may lead to crack initiation on the blade, reducing its useful service life.
Improvement is thus desirable.
In one aspect, the disclosure describes an internally cooled turbine blade comprising an internally cooled turbine blade comprising: an airfoil having a concave side wall and a convex side wall extending spanwise between a platform and a blade tip, and chordwise between a leading edge and a trailing edge, an internal cooling passage within the airfoil extending between a cooling air inlet and a plurality of air outlets, the internal cooling passage including: a serpentine passage having in series a leading edge channel, an intermediate channel and a trailing edge channel, the leading edge channel and the intermediate channel separated by a first dividing wall, the intermediate channel and the trailing edge channel separated by a second dividing wall; and wherein the intermediate channel has an intermediate wall continuously joining the concave and convex side walls, the intermediate wall extending radially between the first and second dividing walls along a central length portion of the intermediate channel for more than half but less than all of the length of the intermediate channel.
In another aspect the disclosure describes a method of reducing creep in an internally cooled turbine blade, comprising: providing a radially extending intermediate wall to continuously join a concave side wall and a convex side wall in an intermediate channel for conveying cooling air through the blade.
Embodiments can include combinations of the above features. Further details of these and other aspects of the subject matter of this application will be apparent from the detailed description included below and the drawings.
Compressed air exits the compressor 5 through a diffuser 6 and is contained within a plenum 7 that surrounds the combustor 8. Fuel is supplied to the combustor 8 through fuel tubes 9 and fuel is mixed with air from the plenum 7 when sprayed through nozzles into the combustor 8 as a fuel air mixture that is ignited. A portion of the compressed air within the plenum 7 is admitted into the combustor 8 through orifices in the side walls to create a cooling air curtain along the combustor walls or is used for cooling to eventually mix with the hot gases from the combustor and pass over the nozzle guide vane 10 and turbine section 11 before exiting the tail of the engine as exhaust.
With reference to
From upstream to downstream, the internal cooling passage begins at the air inlet 22 directing cooling air in a radially outward direction into the leading edge channel 23. A first reverse bend 24 directs the air flow in a radially inward direction into an intermediate channel 25. A second reverse bend 26 directs the air flow in a radially outward direction from the intermediate channel 25 into a trailing channel 27. The air flow passes from the trailing channel 27 and past a series of posts 28 that join the concave side wall 19 and the convex side wall 20 and exhausting from the airfoil 15 through the air outlets 21.
To complete the explanation of the drawings, the blade tip 16 has a blade tip recess 29 that is also supplied with compressed air from the first reverse bend 24 through two ports 30. However the present description is directed to the serpentine internal passage (22, 23, 24, 25, 26, 27, 21) and the internal dividing walls that define it.
The leading edge channel 23 is defined between the leading edge 17 and a first dividing wall 31. The first dividing wall 31 extends radially outwardly from the blade root 13 to the first reverse bend 24 and mechanically joins the concave side wall 19 and the convex side wall 20 of the airfoil 15.
The intermediate channel 25 is defined between the first dividing wall 31 and a second dividing wall 34.
To reduce air flow friction losses, the internal surfaces of the channels 23, 25, 27 are rounded. The leading edge channel 23 and intermediate channel 25 merge arcuately with the first reverse bend 24. The intermediate channel 25 and trailing channel 27 merge arcuately with the second reverse bend 26.
The intermediate channel 25 includes an intermediate dividing wall 36 extending radially parallel to the first and second dividing walls 31, 34.
The localized portions of the concave side wall 19 and the convex side wall 20 that are adjacent to and define the intermediate channel 25 may be susceptible to localized material creep deformation as a result of high stress and high temperature over extended time periods of operation in this particular area. The inventors have provided a method of reducing creep in the internally cooled turbine blade 12 using the radially extending intermediate wall 36 to continuously join the concave side wall and the convex side wall in the intermediate channel 25. The intermediate wall 36 provides a local structural reinforcement that reduces stress in the adjacent local area by distributing stress from a highly stressed area to areas of lower stress. As a result the creep risk is lowered because the stress level is lowered in a susceptible local area.
Referring to
In the example shown, the intermediate wall 36 has a width in a chord-wise direction that is no greater than a minimum width of the first dividing wall 31 and/or the second dividing wall 34. The inclusion of the intermediate dividing wall 36 may enable the reduction of the other dividing walls 31, 34 and hence a variation in weight distribution in the airfoil 15.
Accordingly a creep reinforced zone “C” (dashed lines in
As indicated in
The local divider wall 36 joins the blade convex and concave airfoil walls 19, 20. This local divider wall 36 redistributes loads within the blade material, to reduce local creep strain/stress and improve blade durability. The local divider wall 36 overall length L, thickness and position within the airfoil 15 are also optimized to reduce its adverse effects on the internal cooling flow. By using the local divider wall 36, other design features such as thicker convex and concave airfoil walls 19, 20 are not required to reduce creep strain/stress. This minimizes the blade weight increase and centrifugally-induced loads within the blade and its supporting hub. The strength of these components does not require to be increased and further benefits can then be obtained in terms of total engine weight, performance and operating cost.
The above description is meant to be exemplary only, and one skilled in the relevant arts will recognize that changes may be made to the embodiments described without departing from the scope of the invention disclosed. The present disclosure may be embodied in other specific forms without departing from the subject matter of the claims. The present disclosure is intended to cover and embrace all suitable changes in technology. Modifications which fall within the scope of the present invention will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the appended claims. Also, the scope of the claims should not be limited by the preferred embodiments set forth in the examples, but should be given the broadest interpretation consistent with the description as a whole.
