Patent Application
20140099194
Gas turbine engines operate by burning fuel and extracting energy from the combusted fuel to generate power. Atmospheric air is drawn into the engine from the environment, where it is compressed in multiple stages to significantly higher pressure and higher temperature. A portion of the compressed air is then mixed with fuel and ignited in the combustor to produce high energy combustion gases. The high energy combustion gases then flow through the turbine section of the engine, which includes a plurality of turbine stages, each stage comprising turbine vanes and turbine blades mounted on a rotor. The high energy combustion gases create a harsh environment, causing oxidation, erosion and corrosion of downstream hardware. The turbine blades extract energy from the high energy combustion gases and turn the turbine shaft on which the rotor is mounted. The shaft may produce mechanical power or may directly generate electricity. A portion of the compressed air is also used to cool components of the turbine engine downstream of the compressor, such as combustor components, turbine components and exhaust components.
A turbine engine includes one or more turbine stage. Each turbine stage includes turbine blades extending outwardly from a turbine disk toward an outer surface, which outer surface is referred to herein as a turbine shroud. The first stage, which is the stage closest to the combustor section of the engine, generally extend the shortest distance in a radial direction away from the turbine disk toward the turbine shroud, and also experience the highest temperatures. In each succeeding stage, the turbine blades extend a greater distance in a radial direction away from the turbine disk and toward the turbine shroud, and experience slightly cooler temperatures as the hot gases of combustion expand as they move axially through the turbine engine.
The interface between the turbine blades and the turbine shroud in each turbine stage ideally form a seal, so that the blades can extract as much energy as possible from the flowing, hot gases. The interface between the blades and the shroud experience the hottest temperatures as the gas flows through a turbine stage. If there is a gap between the blades and the turbine shroud, hot gases can escape between the blades and the shroud, resulting in turbine inefficiency. Thus, it is imperative that any gap between the blades and the shroud be minimized if not eliminated. In addition, as the blades rotate at high speeds and high temperatures, the blades will grow from thermal expansion and also from creep, so that the blades tend to wear into the shrouds over a period of time, which assists in maintaining the seal.
Because the sealing surface of the shrouds experience high temperatures, from the hot, oxidative and corrosive combustion gases, as well as abrasion from the rotating blades, it is important to construct the shrouds from high temperature materials that are strong at elevated temperatures, that are corrosion resistant, oxidation resistant, and that also exhibit wear resistance. Depending upon the turbine engine design one or more of the stages may require a high temperature shroud surface that can survive the harsh conditions in the turbine stages.
Turbine shroud materials, particularly in the high pressure turbine stages closest to the combustor, are typically manufactured from materials that have the aforesaid material characteristics. Such materials are expensive and usually are superalloys, such as nickel-based superalloys, iron-based superalloys and cobalt-based superalloys. These shrouds have been constructed both as single pieces and as multi-piece shroud segments. The turbine shroud also includes supporting structure adjacent to its sealing surface which does not see temperatures as high as the sealing surface. These surfaces are out of the gas flow path and so are not constantly exposed to the hot corrosive combustion gases, but these support surfaces, being part of the shroud, also comprise superalloy material.
What is needed is a turbine shroud comprising a plurality of materials in which only the sealing surface comprises a superalloy, while support structure comprises materials that can withstand lower temperatures and lower oxidation and corrosion requirements experienced away from the hot flow path.
A bimetallic ring for use as a turbine shroud in a gas turbine engine is set forth herein. The bimetallic ring forms a sealing surface as a hot gas flow path boundary in the engine. The ring is comprised of two materials. The first material comprises a first portion which is the hot gas flow path sealing surface. The second material comprises a second portion that may be at least a pair of supporting side plates. A dissimilar weld joint joins the sealing surface to the second portion, the at least pair of supporting side plates.
The first material forming the sealing surface further comprises a wrought, oxidation resistant metal alloy having survivability at the hot gas flow path temperatures as the hot gas impinges upon sealing surface. The second material, which is a different material from the first portion and which is out of the hot gas flow path, comprises a material that acts as structural load support for the ring at moderate temperatures. The dissimilar metal weld must be compatible with the first material and the second material. While the dissimilar metal weld is out of the gas flow path, it must provide structural load support at moderate temperatures.
A method for fabricating a bimetallic ring for use as a turbine shroud gas flow path sealing surface in a gas turbine engine is set forth herein. The method comprises the steps of providing a first material, which will form a boundary on which hot gases of combustion will impinge. Because the gases in the hot flow path are hot gases of combustion, the first material is an oxidation resistant metal alloy having survivability at hot gas flow path temperatures. The material is formed into a first portion having a preselected geometry
The method also requires providing a second material. The second material does not experience gas impingement of hot flow path gases. The second material has sufficient strength to provide structural load support for the metallic ring at moderate temperatures. Moderate temperatures as used herein are temperatures away from the hot flow path that are lower than hot gas flow temperatures. The second material is formed into a second portion having a preselected geometry.
The process includes shaping the first material forming the first portion into its preselected geometry and shaping the second material forming the second portion, which may be at least a pair of second plates, into its preselected geometry. Each of the portions has about the same length. The portion is welded to the first portion using a dissimilar weld joint at a junction or joint formed between the second portion and the first portion to form a welded structure. The welded structure may be further worked as required to form an arcuate sealing surface with a pair of flanges, the flanges extending in a substantially transverse direction away from the arcuate sealing surface so that the flanges are not in contact with gases flowing in the hot gas path. The sealing surface has a predetermined radius, which will vary dependent upon engine design, larger engines have a larger radius than smaller engines, which will have a sharper radius of curvature.
Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying drawings which illustrate, by way of example, the principles of the invention.
A gas shroud for use as a sealing surface in a gas turbine engine is set forth herein. The gas shroud interfaces with a rotating blade to form a gas seal. The gas shroud is a metallic ring extending 360° around an engine gas flow path and it may be a single unitary piece formed by forging or welding. Alternatively the gas shroud may be a plurality of arcuate shrouds circumscribing a portion of the circumference of the engine gas flow path, such that when assembled together, forms a metallic ring extending 360° around the engine flow path.
Referring now to
Each of the turbine disks 44, 54 and 54 is mounted on a shaft 20. As hot gases of combustion exit each of stage 1 nozzle 42, stage 2 nozzle 52 and stage 3 nozzle 62, the hot gases striking turbine blades 46, 56, 66 causing precision balanced engine to rotate at high speeds. The hot gases of combustion will contact each of stage 1 shroud 48, stage 2 shroud 58 and stage 3 shroud 68 as the gas traverses turbine section 30 to engine exhaust (not shown) aft of stage 3 blades 66. If there are any gaps between the turbine blades and their respective shrouds, the gas will escape around the gaps, resulting in a loss of efficiency. Efforts are made to maintain the gaps at a minimum to maintain efficiency.
It will be understood by those skilled in the art that a gas turbine engine may have fewer stage or more stages than shown in
Even though each of the shrouds of the present invention have different configurations, each of the shrouds 48, 58 and 68 include common elements. Referring now to
The welded structure can be formed into a shroud for use as stage 1 shroud 48, a stage 2 shroud 58, a stage 3 shroud 68 or any higher stage shroud as required by the engine design by any one of a number of processes. The shroud can be manufactured and formed into a single piece for installation into an engine. The top portion 82 can be formed of a high temperature superalloy such as a nickel-based superalloy, a cobalt-based superalloy, an iron-based superalloy and combinations thereof. While any high temperature superalloy may be used, preferred superalloys include high nickel content, high chromium content and include elements that enable γ′ precipitation strengthening mechanisms, where γ′ is a precipitate having an FCC crystal structure of the form A3B, where A usually is Ni, Co and combinations thereof, and B is Al, Ti and combinations thereof. Those skilled in the art will recognize that γ′ can be formed of other elements (A may include Cr, Mo, V for example), which depends on the overall composition of the alloy selected. Such preferred alloys include Haynes 230, HR-120, Haynes 188, Haynes 25 and INCO® 625. As should be obvious to those skilled in the art, the materials used top portion 82 in stage 1 shroud 48, stage 2, shroud 58 and stage 3 shroud 68 may be different superalloy materials, as the temperature of the hot gases of combustion decreases as the hot gases of combustion expand and move to the exhaust. Clearly, stage 1 shroud 48, which experiences the highest temperatures, must survive the harshest conditions. Top portion 82 can be provided as a wrought material that is rolled or forged, providing an advantage over cast shrouds. Wrought materials allow the grain structure to be controlled so at to take advantage of oriented grains. As an example, the grains in a wrought material can be controlled so that the grains are preferentially elongated in a circumferential direction when the top portion is installed as the sealing surface in the gas turbine engine. Elongation of grains in the circumferential direction improves the erosion resistance of the sealing surface. Although wrought materials are more expensive than cast materials, because the microstructure of a wrought material can be controlled to provide superior mechanical properties, top portion as a wrought material can be with a thinner section in the radial direction than a cast section, with the accompanying advantage of reduced weight.
Alternatively, instead of a single ring, top portion 82 may be fabricated as a plurality of shroud segments that can be joined to form a single ring. The shroud segments can be provided as wrought material, as discussed previously. The wrought material can be provided as a flat plate or the wrought material can be provided as an arcuate shape for subsequent processing.
A pair of side portions 82 can be formed of a moderate temperature material which is less expensive than the high temperature superalloy used to form top portion 82. Since the side portions are assembled to turbine case 16 and support the shroud in the engine, the side portions should have moderate strength at elevated temperatures. Referring again to
Alternatively, when, top portion 82 is fabricated as a plurality of shroud segments that can be joined to form a single ring, side portions also are fabricated as segments that can be joined to top portion 82. Side portions 84 can be provided as wrought material or as cast material, as discussed previously. However, each of side portions should have the same shape as top portion 82 and should be about the same length. When top portion 82 is provided as a flat plate, then side portions 84 should be provided as flat plates as well. When top portion 82 is provided as an arcuate shape, then side portions 84 should be provided as arcuate shapes so that side portions 84 are assembled over top portion 82 such that an inner concave surface of each top portion 84 will mate with opposite sides of outer surface (convex surface) of top portion 82.
Ideally, prior to assembly of side portions 84 to top portion 82, a weld preparation (prep) can be formed on the interfacing surfaces. Thus, for example, when the top portion 82 and side portions 84 are provided as arcuate shapes, a weld prep can be formed on the edges of each side of outer surface (convex surface) of top portion 82 and a weld prep can be formed on the inner concave surface of side portions 84.
Once side portions 84 are fit up to top portion 82, a full penetration weld may be formed to form a welded structure. While the top portion 82 and side portions 84 may be provided so that the weld joint may be made anywhere along the surfaces extending away from the sealing surfaces, top portion and side portions 84 are provided for any particular design to minimize the amount of material provided as top portion 82 in order to minimize expense while maintaining engineering requirements. Because the materials forming the top portion 82 and side portions 84 are different materials, the full penetration weld necessarily is a dissimilar metal weld. The dissimilar metal weld may be accomplished by any technique for full penetration dissimilar metal welds, including but not limited to electron beam welding (EBW), gas tungsten arc welding (GTAW) and gas metal arc welding (GMAW). Welding parameters will depend on the materials used for the top portion 82 and side portions 84. For example, when low alloy steels of grade 22 or grade 91 are utilized with EBW, the fill metal will usually be a shim of Hastelloy® W having a thickness of about 0.020-0.030 inches. When the welding is done using GTAW or GMAW, the filler metal will usually be INCO® 625, except when the base materials include low alloy steels of grade 22 or grade 91, in which case the filler metal will be Hastelloy® W or EPRI P87. However, once the materials are determined, the welding parameters for the dissimilar metal weld should be known to those skilled in the art.
Stress relief of the weld joint also well depend on the materials used for the top portion 82 and side portions 84. However, once the materials are determined, the stress relief heat treatment, if required, for the dissimilar metal weld should be known to those skilled in the art to relieve stresses in the weld and in the heat affected zone (HAZ). Depending upon the materials selected, the stress relief may be of the entire welded structure or it may be a localized stress relief affecting only the weld joint and the heat affected zone.
Each of top portion 82 and side portions 84 may be rough machined or final machined before welding. However, it is preferred that one or both of top portion and side portions 84 only be rough machined before welding.
Preferably, after welding and weld stress relief if required, the γ′ structure may be developed in the seal surface of turbine shroud, formerly the top portion 82. This γ′ structure may be developed before weld stress relief, particularly if the stress relief operation is confined to a local stress relief of the weld and the HAZ, and it may also be developed after final machining. However, developing the γ′ structure after final machining could result in distortion after the precipitation hardening heat treatment.
Final machining preferably is performed on the welded structure after all heat treatment operations.
While the invention has been described with reference to a preferred embodiment, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this invention, but that the invention will include all embodiments falling within the scope of the appended claims.
