Gas turbine engine components with blade tip cooling

A turbine rotor blade for a turbine section of an engine is provided. The rotor blade includes a platform and an airfoil extending from the platform into a mainstream gas path of the turbine section. The airfoil includes a pressure side wall, a suction side wall joined to the pressure side wall at a leading edge and a trailing edge, and a tip cap extending between the suction side wall and the pressure side wall. The rotor blade further includes an internal cooling circuit having a tip cap passage configured to deliver cooling air to the tip cap and a flow accelerator positioned within the tip cap passage of the internal cooling circuit.

TECHNICAL FIELD

The present invention generally relates to gas turbine engines, and more particularly relates to turbine components of gas turbine engines with improved cooling characteristics.

BACKGROUND

Gas turbine engines are generally used in a wide range of applications, such as aircraft engines and auxiliary power units. In a gas turbine engine, air is compressed in a compressor, and mixed with fuel and ignited in a combustor to generate hot combustion gases, which flow downstream into a turbine section. In a typical configuration, the turbine section includes rows of airfoils, such as stator vanes and rotor blades, disposed in an alternating sequence along the axial length of a generally annular hot gas flow path. The rotor blades are mounted at the periphery of one or more rotor disks that are coupled in turn to a main engine shaft. Hot combustion gases are delivered from the engine combustor to the annular hot gas flow path, thus resulting in rotary driving of the rotor disks to provide an engine output.

Due to the high temperatures in many gas turbine engine applications, it is desirable to regulate the operating temperature of certain engine components, particularly those within the mainstream hot gas flow path in order to prevent overheating and potential mechanical issues attributable thereto. As such, it is desirable to cool the rotor blades and stator vanes to prevent or reduce adverse impact and extend useful life. Mechanisms for cooling turbine rotor blades include ducting cooling air through internal passages and then venting the cooling air through holes formed in the airfoil. Internal and film cooling techniques attempt to maintain temperatures that are suitable for material and stress level. However, given the high temperature of engine operation, cooling remains a challenge, particularly in areas such as the turbine blade tips.

Accordingly, it is desirable to provide gas turbine engines with improved blade tip cooling. Furthermore, other desirable features and characteristics of the present invention will become apparent from the subsequent detailed description of the invention and the appended claims, taken in conjunction with the accompanying drawings and this background of the invention.

BRIEF SUMMARY

In accordance with an exemplary embodiment, a turbine rotor blade for a turbine section of an engine is provided. The rotor blade includes a platform and an airfoil extending from the platform into a mainstream gas path of the turbine section. The airfoil includes a pressure side wall, a suction side wall joined to the pressure side wall at a leading edge and a trailing edge, and a tip cap extending between the suction side wall and the pressure side wall. The rotor blade further includes an internal cooling circuit having a tip cap passage configured to deliver cooling air to the tip cap and a flow accelerator positioned within the tip cap passage of the internal cooling circuit.

In accordance with an exemplary embodiment, a gas turbine engine includes a compressor section configured to receive compressed air; a combustion section coupled to the compressor section and configured to receive the compressed air, mix the compressed air with fuel, and ignite the compressed air and fuel mixture to produce combustion gases; and a turbine section coupled to the combustion section and configured to receive the combustion gases. The turbine section defines a combustion gas path and includes a turbine rotor positioned within the combustion gas path. The turbine rotor includes airfoil side walls; a tip cap extending between the airfoil side walls; an internal cooling circuit at least partially defined by the airfoil side walls and including a tip cap passage configured to direct cooling air to the tip cap; and a flow accelerator positioned within the tip cap passage.

DETAILED DESCRIPTION

Broadly, exemplary embodiments discussed herein include gas turbine engines with turbine components having improved cooling characteristics. In particular, exemplary embodiments have turbine blade airfoils with tip cap cooling circuits. The tip cap cooling circuits include one or more flow accelerators to accelerate cooling flow at the tip cap. As a result, cooling may be improved due to increased convection from the high velocity flow and/or increased conduction from enhanced heat flow paths. The increased velocity may also enable more effective use of film cooling holes.

FIG. 1is a cross-sectional view of a gas turbine engine100according to an exemplary embodiment. AlthoughFIG. 1depicts a turbofan engine, in general, exemplary embodiments discussed herein may be applicable to any type of engine, including turboshaft engines. The gas turbine engine100may form part of, for example, an auxiliary power unit for an aircraft or a propulsion system for an aircraft. The gas turbine engine100has an overall construction and operation that is generally understood by persons skilled in the art. The gas turbine engine100may be disposed in an engine case101and may include a fan section120, a compressor section130, a combustion section140, a turbine section150, and an exhaust section160. The fan section120may include a fan, which draws in and accelerates air. A fraction of the accelerated air from the fan section120is directed through a bypass section170to provide a forward thrust. The remaining fraction of air exhausted from the fan is directed into the compressor section130.

The compressor section130may include a series of compressors that raise the pressure of the air directed into it from the fan section120. The compressors may direct the compressed air into the combustion section140. In the combustion section140, the high pressure air is mixed with fuel and combusted. The combusted air is then directed into the turbine section150. As described in further detail below, the turbine section150may include a series of rotor and stator assemblies disposed in axial flow series. The combusted air from the combustion section140expands through the rotor and stator assemblies and causes the rotor assemblies to rotate a main engine shaft for energy extraction. The air is then exhausted through a propulsion nozzle disposed in the exhaust section160to provide additional forward thrust.

FIG. 2is a partial cross-sectional side view of a turbine section of an engine, such as the turbine section150of engine100ofFIG. 1in accordance with an exemplary embodiment. The turbine section150includes a turbine stator200and a turbine rotor250surrounded by a shroud210defining a gas flow path through which hot, combusted air from an upstream compressor section (e.g. combustion section140ofFIG. 1) is directed. Although only one turbine stator200and one turbine rotor250are shown, such stators200and rotors250are typically arranged in alternating axially spaced, circumferential rows. As used herein, the term “axial” refers to a direction generally parallel to the engine centerline, while the term “radial” refers to a direction generally perpendicular to the engine centerline.

The rotor250generally includes rotor blades260(one of which is shown) mounted on a rotor disc (not shown), which in turn is coupled to an engine shaft (not shown). The turbine stator200directs the air toward the turbine rotor250. The air impinges upon rotor blades260of the turbine rotor250, thereby driving the turbine rotor250for power extraction. To allow the turbine section150to operate at desirable elevated temperatures, certain components are cooled. For example, the rotor blades260may be cooled, as described in greater detail below.

FIG. 3illustrates an exemplary aircraft jet engine turbine rotor blade, such as rotor blade260ofFIG. 2, removed from a turbine section.FIG. 3depicts one exemplary embodiment, and other exemplary embodiments may have alternate configurations or arrangements.

The rotor blade260includes an airfoil310, a platform350and a root360. The platform350is configured to radially contain turbine airflow within a shroud (e.g., shroud210ofFIG. 2). The root360extends from the underside of the platform and is configured to couple the blade260to a turbine rotor disc (not shown). In general, the rotor blade260may be made from any suitable material, including high heat and high stress resistant aerospace alloys, such as nickel based alloys, Rene 88, Mar-M-247, single crystal materials, steels, titanium alloys or the like.

The airfoil310projects radially outwardly from the platform350. The airfoil310has two side (or outer) walls312,314each having outer surfaces that together define an airfoil shape. The first side wall312defines a pressure side with a generally concave shape, and the second side wall314defines a suction side with a generally convex shape. In a chordwise direction, the airfoil side walls312,314are joined at a leading edge316and trailing edge318. As used herein, the term “chordwise” refers to a generally longitudinal dimension along the airfoil from leading edge to trailing edge, typically curved for air flow characteristics. The trailing edge318includes trailing edge slots382, discussed below.

In a radial direction, the airfoil side walls312,314extend from a base324at the platform350to a blade tip320. In general, the blade tip320is positioned to rotate in close proximity to the shroud210(FIG. 1) in order to maximize energy extraction. The blade tip320is formed by a tip cap330and squealer tip extensions332. The tip cap330extends between the side walls312,314, typically from leading edge316to trailing edge318. In some exemplary embodiments, the tip cap330is recessed relative to the squealer tip extensions332, which are formed by side walls312,314extending radially beyond the tip cap330. The tip cap330and squealer tip extensions332may be designed to minimize the leakage of hot gasses over the blade tip320of the rotor blade260.

As noted above, the rotor blade260, particularly the airfoil310, is subject to extremely high temperatures resulting from high velocity hot gases ducted from the combustion section140(FIG. 2). If unaddressed, the extreme heat may affect the useful life of an airfoil and/or impact the maximum operating temperature of the engine. As such, cooling is provided for the airfoil310to maintain blade temperature at an acceptable level, as described in greater detail below. Such cooling may include an internal cooling system that directs cooling air from inlets in the root360through internal cavities and passages to cool the airfoil310via convection and conduction. The air flowing through the internal cooling system may flow out of the airfoil310through the trailing edge slots382to provide temperature control of the trailing edge318. Additionally, the cooling air flowing through the internal cooling system may also be supplied to film cooling holes380arranged to provide a cooling film of fluid onto the surface of the airfoil310. InFIG. 3, the film cooling holes380are positioned on the blade tip320, although film cooling holes380may be provided in other locations, such as in the area of the leading edge316and areas immediately aft of the leading edge316.

FIG. 4is a partial cross-sectional view of the rotor blade260ofFIG. 3in accordance with an exemplary embodiment. As shown, the cross-sectional view may generally correspond to a cross-sectional view through a radial-chordwise plane. As discussed above with reference toFIG. 3, the rotor blade260includes the platform350and the airfoil310with leading edge316and trailing edge318. The blade tip320includes the tip cap330and squealer tip extensions332.

FIG. 4particularly shows the interior structure of the rotor blade260, which includes a first cooling circuit400and a second cooling circuit450. Each of the cooling circuits400,450include a portion402,452extending through the platform350and a portion404,454extending through the airfoil310. The platform portions402,452receive cooling air from passages in the root360(FIG. 3) and/or rotor discs (not shown). Such cooling air may be obtained as bleed flow from the compressor section130(FIG. 1). The platform portions402,452respectively deliver the cooling air to the airfoil portions404,454to cool the airfoil310. As described below, the airfoil portions404,454are formed by the side walls312,314and internal structures that direct the air flow through the airfoil310.

The airfoil portion404of the first cooling circuit400includes a first passage410extending from the platform portion402in a generally radial direction, as shown inFIG. 4. In the depicted exemplary embodiment, the first passage410is adjacent to the leading edge316, although other embodiments may have alternate configurations. The airfoil portion404of the first cooling circuit400further includes a second passage420fluidly coupled to the first passage410and extending in a generally chordwise direction. The second passage420generally extends directly adjacent to the tip cap330. In other words, the second passage420may be completely or partially formed by the tip cap330on one side and an interior wall, such as interior wall422, on the other side. As such, cooling air flows through the first passage410to the second passage420where the cooling air flows along the underside of the tip cap330to cool the blade tip320. As described in greater detail below, a portion of the cooling air in the second passage420may flow through film cooling holes in the side walls (e.g., film cooling holes380in side walls312or314as shown inFIG. 3). Any remaining portion of cooling air may flow through the length of the second passage420and out of the first cooling circuit400at outlet428at the trailing edge318, e.g., in a chordwise direction from leading edge316to trailing edge318, although other embodiments may include a second passage that directs cooling air in a different direction along the tip cap. In other configurations, the remaining portion of cooling air may flow from the second passage420into another internal structure of the airfoil portion and/or out of an alternative outlet.

The first cooling circuit400may further include one or more flow accelerators430positioned within the second passage420to accelerate air flow there through. Exemplary flow accelerators430are discussed in greater detail below. Since the first cooling circuit400is configured to deliver cooling air to the blade tip320, the first cooling circuit400may be referred to as a blade tip cooing circuit400.

Now referring to the second cooling circuit450, the airfoil portion454includes a passage460extending from the platform portion452, initially in a generally radial direction and transitioning into a serpentine configuration through the airfoil310, as shown inFIG. 4, although other embodiments may have alternate configurations. The passage460is fluidly coupled to the cooling slots382at the trailing edge of the322of the airfoil310. As such, cooling air flows through the passage460and exits the airfoil310at the cooling slots382to thereby cool the side walls312,314(FIG. 3) of the airfoil310and/or the trailing edge318. A portion of the cooling air flowing through the second cooling circuit450may be provided to film cooling holes (not shown) formed in the side walls312,314(FIG. 3). The second cooling circuit450may be referred to below as a main cooling circuit450.

Although the first cooling circuit400and the second cooling circuit450are described as separate circuits, the first cooling circuit400and second cooling circuits450may be integrated with one another or otherwise be in flow communication. For example, the cooling circuit450may additionally or alternatively deliver cooling air to cool the tip cap330. In one exemplary embodiment, the rotor blade may be a high efficiency, multi-walled turbine blade that is fed from the center body.

FIG. 5is a cross-sectional view of the rotor blade260through line5-5ofFIG. 4. The view ofFIG. 5may be considered a cross-sectional view in an axial-chordwise plane through the second passage420of the first cooling circuit400described with respect toFIG. 4. As such, additionally referring toFIG. 4, the rotor blade260depicted inFIG. 5includes the side walls312,314, the leading edge316, and the trailing edge318.

As also described above, the second passage420receives cooling air (indicated by arrows500,510) from the first passage410and the cooling air flows in a generally chordwise direction. As particularly shown inFIG. 5, the flow accelerator430is positioned in the second passage420to accelerate the cooling flow through the second passage420.

FIG. 6is a cross-sectional view of the rotor blade260through line6-6ofFIG. 4. The view ofFIG. 6may be considered a cross-sectional view in an axial-radial plane through the second passage420of the first cooling circuit400described with respect toFIG. 4. As such, additionally referring toFIGS. 4 and 5, the portion of rotor blade260depicted inFIG. 6includes the side walls312,314, the interior wall422, and the blade tip320, including the tip cap330and the squealer tip extensions332. As particularly shown inFIG. 6, the flow accelerator430is positioned in the second passage420to accelerate the cooling flow through the second passage420, as will now be discussed in greater detail.

Referring now toFIGS. 4-6, the flow accelerator430is positioned in the second passage420to enhance cooling at the blade tip320. The flow accelerator430blocks a portion of the second passage420and the resulting reduction in flow area functions to increase the speed of the cooling flow. The flow accelerator430may have any suitable shape that prevents or mitigates flow or pressure losses through the second passage420. In other words, the flow accelerator430is typically shaped such that the cooling flow smoothly flows around the flow accelerator430, thereby preventing or mitigating pressure losses that may otherwise counteract the acceleration of the cooling flow resulting from the reduction in flow area. As particularly shown inFIG. 5, the flow accelerator430has a tear-drop or airfoil shape to facilitate this function. For example, the flow accelerator430may have a generally curved leading edge that smoothly transitions into generally curved side edges and terminates with a trailing edge. The edges are shaped so as to prevent or mitigate vortices or turbulence through the second passage420. Other shapes that reduce flow area while preventing or mitigating pressure losses may be provided.

The flow accelerator430may have a volume or cross-sectional areas that advantageously accelerate the cooling flows. For example, in the exemplary volume depicted byFIGS. 4 and 5, the flow accelerator430may occupy approximately 50% or more of the volume of the second passage420(e.g., the volume defined by the side walls312,314, the leading edge316, the trailing edge318, the interior wall422, and the tip cap330). As another example, in the exemplary cross-section depicted byFIG. 6, the flow accelerator430occupies approximately 50% of more of the respective cross-sectional area of the second passage420(e.g., the cross-sectional area defined by the side walls312,314, the interior wall422, and the tip cap330). Other volume or cross-sectional area ratios of the flow accelerator430relative to the second passage420may be provided.

As shown in the depicted exemplary embodiment ofFIGS. 5 and 6, the flow accelerator430extends the entire radial height of the second passage420, e.g., from the interior wall422to the tip cap330. As such, the second passage420is separated into a first passage portion424and a second passage portion426that extend on either side of the flow accelerator430.

The configuration of the blade tip320facilitates cooling in a number of ways. For example, the cooling air flowing through the second passage420removes heat from the blade tip320. The flow accelerator430particularly supports this function. Referring toFIG. 5, the airfoil shape of the turbine blade and the position of the flow accelerator430within the second passage420results in a larger cross-sectional area in the forward portion of the second passage420and a smaller cross-sectional area in the aft portion of the second passage420. As such, the velocity of the cooling air through the second passage420is increased as a result of the flow accelerator430. Since velocity of the cooling air is proportional to the cross-sectional area and since heat transfer is proportional to passage velocity, reduced cross-sectional area from the flow accelerator430results in increased heat removal from the tip cap330and side walls312,314via convection.

Additionally, the structure of the flow accelerator430improves heat removal by enhancing the conduction paths through the blade tip320. For example, since the flow accelerator430is within the cooling flow path, the flow accelerator430will generally be at a lower temperature than other portions of the blade tip320, particularly the squealer tip extensions332and tip cap330. As a result of the temperature difference, the heat from the squealer tip extensions332and tip cap330will migrate to the flow accelerator430, thereby increasing the ability of the cooling air flow to remove the heat. Similar heat transfer paths may exist between the squealer tip extensions332and tip cap330to the side walls312,314and/or the interior wall422, thereby decreasing the temperature of the blade tip320. In the exemplary embodiment depicted inFIG. 6, the flow accelerator430extends from the tip cap330to the interior wall422, which may enhance these conduction flow paths.

As an additional cooling mechanism, the flow accelerator430increases the flow velocity while mitigating or preventing flow losses, thereby maximizing the pressure within the second passage420(e.g., relative to utilizing other heat transfer mechanisms such as turbulators or pin fins). The pressure of the cooling air within the second passage420enables a portion of the cooling flow to flow out of the film cooling holes380, thereby enhancing film cooling on the outer surface of the rotor blade260. In conventional turbine blades, loss of pressure within the second passage may adversely impact or prevent film cooling.

FIGS. 7-10depict alternative embodiments of flow accelerators within the second passage of the turbine blade. Generally, unless otherwise noted, the configuration and structure of the turbine blades ofFIGS. 7-10are similar to the turbine blade depicted inFIGS. 2-6.FIGS. 7-10are discussed in greater detail below.

FIG. 7a cross-sectional view of a rotor blade760similar to the view ofFIG. 6. The view ofFIG. 7may be considered a cross-sectional view in an axial-radial plane through a tip cooling passage720of a cooling circuit. As such, the portion of rotor blade760depicted inFIG. 7includes the side walls712,714, the interior wall722, and the blade tip724, including the tip cap740and the squealer tip extensions726. As particularly shown inFIG. 7, the flow accelerator730is positioned in the passage720to accelerate the cooling flow through the passage720. As compared to the flow accelerator430ofFIG. 6, the flow accelerator730extends from the tip cap740to only a portion of the radial distance (or radial height) between the tip cap740and the interior wall722. For example, the flow accelerator730extends approximately three-quarters of the radial distance, although other dimensions may be provided.

FIG. 8is a cross-sectional view of a rotor blade860similar to the view ofFIG. 5discussed above. LikeFIG. 8, the view ofFIG. 5may be considered a cross-sectional view in an axial-chordwise plane through a tip cooling passage820of a cooling circuit. As above, the passage820receives cooling air from within the rotor blade860and the cooling air flows in a generally chordwise direction.

FIG. 8additionally includes a number of flow accelerators830,832,834positioned in the passage820to accelerate the cooling flow through the passage820. The flow accelerators830,832,834are generally airfoil-shaped and may be arranged in any suitable manner to accelerate cooling flow while preventing or mitigating pressure loss.

FIG. 9is a cross-sectional view of a rotor blade960similar to the view ofFIG. 5discussed above. LikeFIG. 9, the view ofFIG. 5may be considered a cross-sectional view in an axial-chordwise plane through a tip cooling passage920of a cooling circuit. As above, the passage920receives cooling air from within the rotor blade960and the cooling air flows in a generally chordwise direction.

FIG. 9additionally includes a number of flow accelerators930in the form of multiple pins positioned in the passage920to accelerate the cooling flow through the passage920. The flow accelerators930are generally cylindrical, oval, or airfoil shaped and may be arranged in any suitable manner to accelerator cooling flow while preventing or mitigating pressure loss.

FIG. 10a cross-sectional view of a rotor blade1060similar to the view ofFIG. 6. The view ofFIG. 10may be considered a cross-sectional view in an axial-radial plane through a tip cooling passage1020of a cooling circuit. As such, the portion of rotor blade1060depicted inFIG. 7includes the side walls1012,1014, the interior wall1022, and the blade tip1016, including the tip cap1042and the squealer tip extensions1040. As particularly shown inFIG. 10, the flow accelerator1030is positioned in the passage1020to accelerate the cooling flow through the passage1020. In this exemplary embodiment, the flow accelerator1030has a radial portion1032that extends from the tip cap1042to the interior wall1022. In an axial cross-section, the radial portion1032may have a shape such as the flow accelerator430described above. The flow accelerator1030further includes a lateral portion1034that extends between the side walls1012,1014(or from a first side of the radial portion1032to side wall1012and from a second side of the radial portion1032to side wall1014). The lateral portion1034may have a hydrofoil or airfoil shape. As such, in the depicted exemplary embodiment, the passage1020may be divided into four segments1036,1037,1038,1039to accelerate cooling flow while preventing or mitigating pressure loss.

FIG. 11is a partial cross-sectional view of a turbine rotor blade1160(e.g., the outer radial portion of the blade1160) in accordance with an alternate exemplary embodiment, andFIG. 12is a partial cross-sectional view of the turbine rotor blade1160along line12-12ofFIG. 11in accordance with an exemplary embodiment. As above, the turbine rotor blade1160may include side walls1162,1164, a leading edge1166, and a trailing edge1168. The turbine rotor blade1160further includes a tip cap1130and squealer tip extension1132.

The turbine rotor blade1160ofFIGS. 11 and 12includes cooling circuits1100,1150similar to the cooling circuits400,450described with respect toFIG. 4, although the cooling circuits1100,1150may be modified as necessary or desired. In the depicted exemplary embodiment, the cooling circuit1100includes a first passage1110that delivers cooling air (e.g., cooling air1112) to a second passage1120, which extends in a chordwise direct such that the cooling air cools the tip cap1130, as described above. As also described above, an accelerator1122may be positioned within the second passage1120to accelerate the cooling air flowing through the second passage1120, thus resulting in improved convection and conduction cooling while mitigating and/or preventing flow losses.

In this exemplary embodiment, the accelerator1122may be hollow or otherwise include an outer wall1124and an interior space1126. As best shown byFIG. 11, the second cooling circuit1150may deliver a cooling flow (e.g., cooling flow1152through an interior or central passage) through an inlet1170in the interior wall1172to the interior portion1126of the accelerator1122. The cooling flow1152may impinge the underside of the wall1124to provide impingement cooling, as well as convection and/or conduction cooling. The cooling flow1152then flows in an aft chordwise direction through the accelerator1122and out of the accelerator1122at an accelerator outlet1128. The cooling flow1152subsequently flows through the second passage1120(e.g., with the cooling flow1112from the first cooling circuit1100) and out of the turbine rotor blade1160at the trailing edge1168. In general, the cooling flow1152flows through the accelerator1122and second passage1120in a low-loss manner, and in some exemplary embodiments, the cooling flow1152may function to increase pressure within the second passage1120, thereby potentially enhancing the overall cooling flow.

Accordingly, turbine rotors with improved blade tip cooling are provided. Exemplary embodiments of the turbine blades discussed above have resulted in an ability to increase engine temperature, thereby improving fuel consumption. In addition to the flow accelerators discussed above, exemplary embodiments may also use turbulators, depressions, and other techniques that may enhance tip cap cooling. In general, the flow accelerators may be formed in a selected pattern or array to provide optimum cooling. Computational fluid dynamic (CFD) analysis can additionally be used to optimize the location and orientation of the flow accelerators and cooling circuits and passages. Exemplary embodiments promote the service life and/or enhanced performance in a cost-effective manner. The turbine blades produced according to exemplary embodiments may find beneficial use in many industries including aerospace, but also including industrial applications such as electricity generation, naval propulsion, pumping sets for gas and oil transmission, aircraft propulsion, automobile engines, and/or stationary power plants.