Patent Application
20090169361
The exemplary embodiments relate generally to gas turbine engine components and more particularly to turbine nozzle segments having improved cooling.
Gas turbine engines typically include a compressor, a combustor, and at least one turbine. The compressor may compress air, which may be mixed with fuel and channeled to the combustor. The mixture may then be ignited for generating hot combustion gases, and the combustion gases may be channeled to the turbine. The turbine may extract energy from the combustion gases for powering the compressor, as well as producing useful work to propel an aircraft in flight or to power a load, such as an electrical generator.
The turbine may include a stator assembly and a rotor assembly. The stator assembly may include a stationary nozzle assembly having a plurality of circumferentially spaced apart airfoils extending radially between inner and outer bands, which define a flow path for channeling combustion gases therethrough. Typically the airfoils and bands are formed into a plurality of segments, which may include one or two spaced apart airfoils radially extending between an inner and an outer band. The segments are joined together to form the nozzle assembly. The band may include one or more flanges for attaching the nozzle assembly to other components of the gas turbine engine.
The rotor assembly may be downstream of the stator assembly and may include a plurality of blades extending radially outward from a disk. Each rotor blade may include an airfoil, which may extend between a platform and a tip. Each rotor blade may also include a root that may extend below the platform and be received in a corresponding slot in the disk. Alternatively, the disk may be a blisk or bladed disk, which may alleviate the need for a root and the airfoil may extend directly from the disk. The rotor assembly may be bounded radially at the tip by a stationary annular shroud. The shrouds and platforms (or disk, in the case of a blisk) define a flow path for channeling the combustion gases therethrough.
As gas temperatures rise due to the demand for increased performance, components may not be able to withstand the increased temperatures. Higher gas temperatures lead to higher metal temperatures, which is a primary contributor to distress. Distress may cause cracking or holes to form within these areas, leading to decreased performance and higher repair costs. Higher pressure and temperature areas suffer the greatest distress. As shown in
In one exemplary embodiment, a turbine nozzle segment may have a band having a flowpath side, a non-flowpath side, a flange extending radially from the non-flowpath side and an aft end. The nozzle segment may further include a plurality of airfoils having trailing edges and extending radially from the flowpath side. The nozzle segment may also have a plurality of cooling holes disposed in the flange, the cooling holes directed at the aft end of the non-flowpath side of the band between the trailing edges.
In another exemplary embodiment, a turbine nozzle assembly may include a plurality of arcuate turbine nozzle segments joined together to form an annular ring, each of the plurality of arcuate segments having a band having a flowpath side, a non-flowpath side, a flange extending radially from the non-flowpath side and an aft end. The nozzle segment may further include a plurality of airfoils having trailing edges and extending radially from the flowpath side. The nozzle segment may also have a plurality of cooling holes disposed in the flange, the cooling holes directed at the aft end of the non-flowpath side of the band between the trailing edges.
As shown in
Airfoils 136 extend radially between the inner band 120 and outer band 122 for directing the flow of combustion gases through the turbine nozzle assembly 116. The airfoils 136 have a leading edge 138 on the forward side of the turbine nozzle segment 118 and a trailing edge 140 on the aft side of the turbine nozzle segment 118. The airfoils 136 may be formed of solid or hollow construction. Hollow airfoils may include one or more internal cooling passages for cooling the airfoil and providing film cooling to the airfoil surfaces. Other hollow airfoils may include one or more cavities for receiving a cooling insert. The cooling insert may have a plurality of cooling holes for impinging on the interior surface of the hollow airfoil before exiting as film cooling through holes in the airfoil. Any configuration of airfoil known in the art may be used.
Band, as used below, may mean the inner band 120, the outer band 122 or each of the inner band 120 and outer band 122. The band may have one or more flanges 132 extending radially from the non-flowpath side 126, 130. At least one of the flanges 132 may be located near the aft side of the nozzle segment 118, such as, but not limited to, flange 134 in
A plurality of cooling holes 144 may be disposed within the flange 134. The cooling holes 144 may have an inlet 146 at the plenum 142 on the upstream side of the flange 134 and an outlet 148 on the downstream side of the flange 134. The inlet 146 may receive cooling air from the plenum 142 and flow the cooling air through to the outlet 148. The cooling hole 144 and outlet 148 may be arranged so that the outlet 148 is directed at the aft end 150 of the band, so as to impinge on the aft end 150. The outlets 148 may have any shaped known in the art. Further, the holes 144 may be formed in any manner known in the art, such as, but not limited to, electrodischarge machining, electrochemical machining, laser drilling, mechanical drilling, or any other similar manner.
In one exemplary embodiment, as shown in
In one exemplary embodiment, a thermal barrier coating (TBC) 160 may be applied to the band flowpath surface 124, 128. The TBC may be between about 5 mils and about 25 mils thick. Any TBC known in the art may be used. In one exemplary embodiment, the TBC may be a three layer TBC having a MCrAlY first layer, where M is selected from the group of Ni and Co, an aluminide second layer, and a yttria-stablized zirconia (YSZ) third layer. In another exemplary embodiment, a two layer TBC may be used where platinum aluminide or aluminide may be used in place of the MCrAlY first layer and the aluminide second layer.
By providing cooling holes in these areas and in particular by impinging cooling air in these areas, the metal temperature may be reduced, leading to less distress and less likelihood of forming a crack or hole. As such, the turbine nozzle segment will last longer leading to less repairs and/or replacements over time for the gas turbine engine.
This written description discloses exemplary embodiments, including the best mode, to enable any person skilled in the art to make and use the exemplary embodiments. The patentable scope is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.
