Patent Application
20080279695
1. Technical Field
This invention relates to the internal cooling of gas turbine engine, turbine airfoils and particularly the end portions thereof.
2. Background Art
Modern gas turbine engines operate at temperatures approaching 3000° F. Accordingly, it is a common practice to cool various components employed in such engines with air provided by the engine's compressor. Perhaps the most critical components to cool with compressor air are the first stage turbine blades and vanes which are exposed to products of combustion exiting the engine's combustor.
It is well known to provide such compressor discharge cooling air to first stage turbine blades and vanes by routing such air through passages internally of the airfoil portions thereof. Such passages may be cast into the airfoil portions or drilled into the blades or vanes by mechanical or electrochemical machining processes.
In the case of turbine blades and vanes for large industrial gas turbine engines, it is a common practice to employ shaped tube electrochemical machining to form cooling air passages which extend radially from the inner end of the airfoil to the outer end thereof. For enhanced convective cooling, the cooling air passages often include discontinuities in the walls thereof to enhance the turbulence of the flow of cooling air through the passages by eliminating the boundary layer of airflow along the passage walls. Such discontinuities, often referred to as turbulence promoters or turbulators, may take the form of grooves or ridges in the cooling passage walls.
While such turbulators enhance the convective cooling of the interiors of turbine blades and vanes, they necessarily increase the losses associated with the flow of cooling air through the passages and thus adversely affect the overall efficiency of the engine. Therefore, it has been the conventional wisdom to use such turbulators only where they are most necessary from the standpoint of thermal loading. It is generally accepted in the prior art that the locations where internal cooling of turbine blades and vanes is most critical (where thermal loads are greatest) are those locations in the blade or vane airfoils intermediate the root and tip portions thereof. Accordingly, as a result of qualitative analyses of the operating characteristics of blades and vanes, it has been the practice to provide such turbulators only in the intermediate portions of the internal cooling passages of turbine blades and vanes, the root and tip ends of the passages being smooth to minimize the inefficiencies associated with the creation of turbulent flow therein.
However, inspections of modern industrial gas turbine engines, as part of the routine overhaul and maintenance thereof, has revealed that the blades and vanes of such engines experience significant and often unanticipated thermal stress at the ends thereof as evidenced by, for example, cracking in the blade shrouds, such as, in the fillet where the shroud joins the blade. Several solutions to such thermal stress and damage to the blade have been proposed and typically involve a rather complex distribution of additional cooling passages and chambers in the shroud. While such cooling schemes have met with limited success, they greatly increase the complexity of the internal cooling passage configuration and thus greatly increase the complexity and manufacturing costs of the blade. These increased costs may more than offset the savings in operating costs associated with having smooth bores at the radially inner and outer ends of the airfoil cooling passages.
The present invention is predicated on the recognition that the qualitative analyses which led to the implementation of turbulators only in the intermediate portions of blade and vane radial cooling passages may have failed to take into account factors which would cause destructive thermal loading at the end portions of the blades and vanes, for example, at blade shrouds through which the unturbulated portions of the cooling passages extend.
One factor which would give rise to destructive thermal loading of the blade and vane end portions is a reduced total airflow through the cooling passages due to anomalies in the cooling air flow circuit beginning with the gas turbine engine's compressor and terminating with the blade or vane itself. Such anomalies include, for example, partial blockage of the flow passages with foreign matter, anomalies in the operation of the engine's compressor, wear of rotating seal components etc.
Another factor which theoretically can cause destructive thermal loading of blade and vane end portions is a deviation from a normal (uniform) temperature profile at the exit of the engine's combustor. Typically, gas turbine engine combustors are designed to provide combustion gases at a generally uniform temperature profile across the flow path of the engine's products of combustion. Foreign matter or pollutants in the engine's fuel system can cause blockage of some of the full nozzles in the combustor, resulting in asymmetries in the temperature profile across the combustor exhaust, thereby resulting in hot spots in the vanes and nozzles. Moreover, when replacement vanes and blades are employed in engines with unknown nominal operating parameters such as combustion exhaust temperature profiles, it would most efficacious to provide such blades with sufficient turbulation at the ends of the cooling passages to accommodate any anomalies in engine operation such as unevenness in the temperature profile at the combustor exhaust.
Recognizing that the heretofore common practice of providing turbulation only at the intermediate or medial portion of blade and vane cooling passages may not provide adequate convective cooling of gas turbine engine blades and vanes, in accordance with the present invention, turbulence promoters are provided in such blades and vanes at the radial extremities thereof. In a preferred embodiment of the present invention, in a turbine blade having radial cooling holes substantially along the entire length thereof, turbulence promoters are provided all the way to the tip of the blade including through any outer shroud thereof. The turbulence promoters may take on any of various known shapes such as annular or partially annular ribs or grooves.
In accordance with another aspect of the present invention, the thermal performance of prior art blades and vanes may be improved upon by adding turbulation promoters to the smooth walled portions of radial cooling channels, thereby restructuring such channels to increase the turbulent flow and thus the convective cooling provided in such smooth walled portions to accommodate the unanticipated destructive thermal loading outlined above.
It has been determined that perhaps counterintuitively, adding such turbulation promoters to such smooth walled portions of the cooling channels does not unacceptably lower the operating efficiency of the associated engine nor does it appreciably increase the manufacturing costs of the blades and vanes since fully turbulated holes may be formed without undue attention to the depth of placement of the tooling which forms the turbulators at the beginning and conclusion of the turbulator forming process.
Finally, it is believed that at least in the case of the provision of turbulators in the radially outer ends of shrouded turbine airfoils, the enhanced convective cooling of the shroud by a resultant turbulent cooling may reduce the need for stress reducing structures such as fillets and the like, thereby minimizing the size and weight of such structures as well as reducing the need for added cooling holes, passages and other fluid handling structural intricacies in the shroud and, in general, increase the overall mechanical and thermal capacity of such blades.
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Turbine blade 10 may be formed from any suitable material known in the art such as a nickel based superalloy. To improve the cooling characteristics of the turbine blade 10, each of the cooling passages 20 has a plurality of turbulation promoters (turbulators) disposed therealong, not only within airfoil portion 15, but also along the radially inner and outer portion thereof, within shrouds 30 and 34.
Referring now to
A plurality of turbulation promoters (turbulators) 90 are incorporated into the passage 20. The turbulation promoters may comprise arcuately shaped trip strips which have a height e and which circumscribe an arc of less than 180 degrees. The ratio of e/D is preferably in the range of from 0.05 to 0.30. Trip strips 95 may be annular or take the form of spaced arcuate members (see
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The pairs of trip strips 95 are preferably aligned so that the gaps g of one pair of trip strips 95 is aligned with the gaps g of adjacent pairs of trip strips 95. It has been found that such an arrangement is desirable from the standpoint of creating turbulence in the flow in the passageway 20 and minimizing the pressure drop of the flow.
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As set forth hereinabove, the cooling passages shown in
While the turbulence promoters are shown and described herein as acute in shape and circumscribing somewhat less than 180 degrees, it will be understood that fully annular turbulence promoters or turbulence promoters of any of various other known shapes such as full or partial helices may be employed with equal efficacy and may be formed by methods other than the aforementioned electrochemical machining operation, such as ordinary mechanical drilling and tapping methods.
Also, while the present invention as shown and described within the context of a blade or vane manufactured in accordance with the present invention, the present invention is equally applicable in the improvement of prior art blades or vanes wherein only the intermediate portions of the cooling air passages are turbulated. In such cases, the smooth bore portions of the cooling air passages may be machined by any of the methods mentioned hereinabove to add turbulence promoters thereto, resulting in the advantages and benefits discussed hereinabove.
Furthermore, while the invention herein has been described in connection with the outer shroud of a gas turbine engine turbine blade, it will be understood that this invention is equally applicable to inner turbine blade shrouds as well as inner or outer vane platforms and shrouds.
Therefore, it will be appreciated that various embodiments and applications of the present invention beyond those specifically discussed and illustrated herein are contemplated and it is intended by the appended claims to cover such embodiments and applications as full within the true spirit and scope of this invention.
