Patent Application
20190024520
The presented inventions relate generally to gas turbine engines, and more specifically, to internal cooling of blades and vanes subjected to high temperatures inside a gas turbine engine. The inventions have particular advantages for improved turbine engine efficiency, performance, and safety.
Gas turbine engines provide propulsion to all jet aircraft, turbo-props, and helicopters, as well as to some marine vessels. Additionally, some gas turbine engines are used to convert natural gas to electricity to supply a significant fraction of U.S. electrical power.
Most gas turbine engines operate in a similar manner, that is, by combusting fuel in a stream of compressed air to extract energy. Essentially, a gas turbine engine has a compressor in fluid communication with a turbine with a combustor therebetween. The compressor and the turbine may have a series of stages, each associated with varying properties such as diameter, blade angle, total pressure, etc. Ambient air is received at an inlet of the compressor before being compressed and discharged into the combustor at a substantially higher pressure and temperature. It is mixed with fuel in the combustor and burned, thereby increasing the temperature and velocity of combustion gases (i.e., gases exiting the combustor). The hot combustion gases are routed through the turbine blades causing them to rotate such that mechanical shaft power may be extracted to drive a shaft, propeller or fan (including the compressor, for example). Excess exhaust pressure can be used to provide thrust.
The efficiency and power output of a gas turbine engine depends on many factors including, but not limited to, the size, pressure and temperature levels achieved. Higher operating temperatures are desired, but material limitations restrict operating temperatures to prevent cracking, softening, or other mechanical failures of engine components. Therefore, improved cooling schemes are desired to allow for higher operating temperatures without material failures.
Gas turbine engine technology is constantly challenged to operate at higher efficiencies, leading to a push towards stoichiometric combustion and ever higher combustor outlet temperatures. In a modern gas turbine engine, temperatures can exceed the vane (e.g., nozzle guide vane) material limits by 600° F. or more, necessitating internal convective cooling and film-cooling schemes in addition to the use of thermal barrier coatings. Internal convective cooling is often inadequate alone, particularly in the stagnation region of the vane (e.g., the area around the stagnation point, where combustion gas flow is bifurcated around the vane and a portion of the flow is at least temporarily brought to rest). Both internal and film-cooling approaches can lead to significant performance penalties in the engine, due to thermal mixing (e.g., cooling air is dumped into the hot stream of combustion gases) and total pressure losses (e.g., siphoning high pressure cooling air from the compressor and dumping it into the core flow without extracting useful work).
In early gas turbine engines, convection cooling was extensively used in a manner in which the rotating blade acted essentially as a single-pass cross-flow heat exchanger. Compressed air would flow radially through internal passages in one-direction from root of a blade to tip, propelled by pressure differences and centrifugal force, thereby removing heat convectively transferred to the blade. Improvements in modern manufacturing technology means that it may now be possible to create a serpentine web of cooling channels within a blade or vane, thereby effectively turning the system into a multi-pass heat exchanger with improved cooling capabilities. Furthermore, cooling air may also be exhausted through small exhaust ports onto the external airfoil surface to form a thin protective layer of cooling air around the vane, hence the name “film-cooling.”
Currently, turbine components are typically cooled by circulating cooling air (i.e., air that is cooler than the combustion gases) from the compressor through internal passages located inside the given component. The circulated cooling air may provide a convective cooling effect as it passes through the internal passages of the component. The cooling air is heated as it cools the component (although it is still cooler than the combustion gases) and may then be discharged through small exhaust ports in an exterior surface of the component to form a film-cooling effect. In a modern gas turbine engine, about 20% of the air from the compressor may be bled off for cooling and sealing related to vanes and turbine blades. For example, vanes may utilize cooling air from the compressor for film-cooling. Additionally, a cooling system may be used to prevent any hot core flow gases from flowing over blade-attachment discs or to control tip clearances between turbine blades and an outer casing (e.g., active clearance control).
Typically, the first stage turbine vanes of a gas turbine engine are exposed to the highest temperatures as they are nearest the combustor on the downstream side. The temperature may then decrease in later turbine stages due to dilution of hot combustion gases with cooling air, relative velocity effects and power extraction (e.g., gas expansion causes a drop in temperature). One challenge associated with cooling these critical vanes is that the pressure differential between the cooling air and the hot combustion gases which flow around the airfoil of the vane needs to be minimal to maintain efficiency. Cooling air for a first stage turbine vane may be received from the compressor and routed internally through the vane before being fed through an exhaust port (e.g., a film-cooling hole) in the vane and into the core flow (the primary path through the engine including the compressor, combustor, and turbine). Combustion gases may also be routed to (or near) the same exhaust port from an external side of the vane in the core flow. In this regard, if the combustion gas pressure were higher than the pressure of cooling air routed from the compressor, the cooling air would fail to exhaust from the vane and hot combustion gases would enter the vane through the exhaust ports. This would be undesirable as it would subject the internal passages of the vane to high temperatures and stall the cooling system, likely leading to material failure caused by overheating. During normal operation of a gas turbine engine, however, the laws of thermodynamics require that, due to combustion inefficiencies, a small pressure drop occurs in the combustor such that cooling air routed from the compressor through a bypass (e.g., around the combustor) maintains a higher pressure than the combustion gases exiting the combustor. This means that the core flow pressure at the first row of vanes in the turbine directly after the combustor must be lower than at the exit of the final stage of the compressor. It is this pressure difference that is used to drive cooling air through internal passages of the vanes. In this respect, improvements in combustor design over the last several years have been both an advantage and a disadvantage for cooling engineers. Improvements in compressor and combustor designs have led to higher compressor pressure ratios and lower combustor pressure losses, respectively, such that more force is available to drive cooling air to cool the hotter aft parts of the engine. On the other hand, with increasing compression ratios the air within the compressor naturally reaches higher exit temperatures (e.g., around 900K prior to combustion) reducing the effect that the cooling air has on the turbine vanes. Furthermore, cooling air is expensive from an efficiency point of view because it would be ideal to “waste” as little air from the compressor as possible for cooling purposes so that it can be preserved for combustion and/or thrust.
Generally, the efficiency of a convective cooling system is measured by the ratio of actual temperature increase of the cooling air to the theoretical maximum temperature increase possible. In this regard, a slight temperature increase in the cooling air indicates low cooling efficiency whereas a temperature increase up to the temperature of the component to be cooled (theoretical maximum) indicates 100% cooling efficiency. In previous methods of cooling a turbine vane which utilize a single impingement point, cooling air is directed to impinge on a surface and then is exhausted. This brief interface between the cooling air and the heated component limits the amount of heating capacity used and thereby limits cooling efficiency.
Accordingly, prior art methods fail to effectively cool vanes and blades of jet turbine engines above current operating temperatures and compromise engine efficiency. It would be advantageous to develop a cooling system and method with improved cooling such that operating temperatures and associated efficiencies could be increased.
The specific power of a gas turbine engine generally increases with inlet temperature. In the last 25 years, however, maximum material temperature limitations have significantly hindered progress toward operating temperature increases. Improved cooling of the engine blades and vanes may allow higher inlet temperatures without stressing materials beyond temperature limitations.
An important metric of cooling turbine vanes and blades is the thermal resistance (a measure of the ease with which the vane can dissipate heat from a surface). As will be appreciated, the thermal resistance of a vane is related to the thermal conductivity of the material forming the vane. However, the thermal resistance is also related to the physical configuration of internal passages in the vane used for cooling and operating conditions. By way of example, thermal resistance of a vane will change based on the design of the internal passages, the thickness of material separating the internal passages from other components of the vane, the flow rate of air through the internal passages, heat load, etc. An embodiment of the presented inventions includes placement of micro-structures in regions of high heat exposure in vanes and blades to improve the thermal resistance thereof.
The disclosed turbine vane cooling concepts provide enhanced internal cooling effectiveness via the use of micro-structured impingement surfaces in the hottest regions of a turbine vane (e.g., stagnation point and suction side heating peak) which may be combined with micro-channel cooling of vane side walls, eliminating the need for film-cooling. Performance degradations due to thermal mixing and total pressure losses can be mitigated by reducing the amount of cooling air introduced into the core flow through the first few turbine stages. This can be achieved by improving internal cooling effectiveness, thereby reducing or eliminating the need for film-cooling, and relying solely, or more heavily, on internal cooling. It should be noted that although higher air flow rates through cooling systems result in better heat transfer, the increase typically reduces engine efficiency (e.g., by bleeding pressurized cooling air from the core flow). Therefore, in conventional systems, high flow rates are common but may result in sacrificing performance and efficiencies. The concepts disclosed herein may reduce the needed air flow rate resulting in improved efficiency over conventional systems. Total pressure losses may also be limited by minimizing perturbations in the combustion gas flow caused by exhausting cooling air into the core flow.
An embodiment of the presented inventions incorporates flow channel cooling techniques in a gas turbine engine vane. Notably, as used herein, a flow channel may refer to any internal passage in a turbine vane which serves as a conduit to move cooling air through a vane within a spanwise cross-sectional plane. Micro-channel, as that term is used herein, generally refers to a flow channel disposed adjacent to a suction-side wall, pressure-side wall, or trailing edge and typically do not contain micro-structures, although in certain embodiments they may. Generally, reference to a flow channel refers to an internal passage containing micro-structures disposed adjacent to the leading edge region.
One form of constructing a vane having a flow channel cooling system includes a plurality of thin plates or foils which may be laminated together to form a block. Often, the foils may be thin metallic strips. Apertures, recesses, and other design features may be cut through or into the foils at various locations as needed. For example, apertures may be cut through the foils at a common location at opposite ends of recessed portions such that, when the stack of foils is laminated, the apertures align to form air distribution plenums though the stack in a spanwise direction (typically normal to the face of each foil) and recessed portions form flow channels within a spanwise cross-sectional plane, connecting air distribution plenums. Additionally or alternatively, apertures may also be cut into some foils and omitted from others such that when stacked, apertures on adjacent foils may be combined to form flow channels of a greater width than just a single foil. Such a flow channel may be capped on each end by a foil or foils having a solid surface in the respective area.
Each of the air distribution plenums may essentially be considered a conduit which passes through at least a portion of the vane to supply (i.e., inlet plenum) or receive (i.e., outlet plenum) cooling air to and from flow channels, respectively. The air distribution plenums may be connected via a plurality of shallow channels (i.e., flow channels) formed from recessed portions, apertures and/or walls formed in or defined by the foils. It should be appreciated that, in some embodiments such internal passages may be formed into the vane after the foils are laminated/bonded. That is, internal passages may be milled or otherwise formed into the vane and in some embodiments, a vane may be 3D printed such that the air distribution plenums, flow channels, and/or micro-structures are formed in each layer of the vane as it is built.
Recessed portions may be etched into at least one face of a foil through mechanical or chemical means. For example, although various depths of recessed portions and various thicknesses of foils are envisaged, the recessed portions may have a shallow dimension on the order of 10-50 microns deep prior to lamination of a 0.002″ thick foil. The air distribution plenums may be attached to a manifold on one or both ends of the vane. The manifolds may be operative to circulate cooling air through the air distribution plenums and the flow channels. Notably, more complex geometries and/or branching networks of internal passages may be utilized and the embodiments described herein are exemplary in nature.
The thicknesses of the foils may be equal or may differ from one foil to another or even within a single foil. For instance, to tailor a surface such as a leading edge to have desired thermal properties (e.g., increased cooling in one area with reduced weight in another), it may be desirable to use differing thicknesses to create micro-structures and flow channels having differing properties. Furthermore, the foils may be of one-piece construction or may be otherwise configured, for example, as multi-piece foils that collectively define the internal passages and other features.
Once the foils are formed, they may be disposed in a stack such that the flow channels are aligned. Each foil may have alignment holes configured to receive an alignment rod during stacking to ensure proper alignment of the flow channels. Once aligned, the mating faces of adjacent foils may be laminated or bonded together. In one embodiment, such lamination may be performed in a diffusion bonding process. In some embodiments, other processes such as brazing or soldering may be utilized.
It is contemplated that vanes in accordance with an embodiment of the presented inventions may be fabricated from Hastelloy X, a high-temperature super alloy. It is noted that any suitable material may be used instead of or in addition to Hastelloy X. Some non-limiting examples include Inconel MA754, Haynes 230, Waspaloy, or Inconel 617. Hastelloy X may be preferable for applications utilizing diffusion bonding manufacturing techniques in order to prevent delamination.
In accordance with an embodiment of the presented inventions, a micro-structured impingement surface may comprise a plurality of fins forming a finned array. One or more of such fins may be etched or cut into one or more foils and may serve to increase the surface area available for convective cooling on a high temperature region of a vane. Numerous (e.g., hundreds or thousands) fins may be disposed on an internal impingement surface of an outer wall of a vane in the leading edge region depending, for example, on the physical dimensions of the vane. As cooling air is circulated over the fins, heat may be transferred from the fins and into the air, thereby cooling the otherwise temperature vulnerable leading edge region(s) of the vane.
In order to maximize cooling for micro-structured impingement surfaces, it may be desirable to utilize a small fin size with a high velocity of impingement cooling air. However, this may cause an undesirable drop in air pressure. Larger fins may cause a reduction in the velocity of cooling air circulating through the impingement region thereby retaining air pressure but may reduce cooling effectiveness due to the lower air velocities. Therefore, a balance may be struck between pressure, velocity, and cooling effectiveness. In other words, cooling effectiveness must be balanced against overall engine efficiencies.
An impingement cooling region of a vane may be segmented into multiple impingement cooling regions. For example, a dual region design may be utilized such that an inside surface of each heating peak (i.e., stagnation point or suction-side heating peak) is disposed within an impingement cooling region (i.e., two regions with flow entering between the peaks and wrapping around toward the sides of the vane in opposite directions). Such a segmented design may promote higher cooling air pressures or velocities and lower cooling air temperatures at the heating peaks as compared to a single channel design. In order to cool the surface of a vane downstream of an impingement region, micro-channels may be incorporated into the vane side walls in areas where fins or other micro-structures are omitted. In this regard, regions of a vane which are not subjected to temperatures as extreme as those near the leading edge may nonetheless be internally cooled by convection. Each foil may comprise a portion of such micro-channels or certain foils (e.g., every other foil, every 6th foil, etc.) may comprise a solid portion in the side wall regions which serves to separate micro-channels disposed in adjacent foils.
In an embodiment, a first inlet plenum may supply cooling air to an impingement region(s) containing micro-structures. Furthermore, to eliminate the need to exhaust cooling air into the core flow, after the cooling air has provided impingement cooling, the cooling circuitry may allow for side wall cooling via a second inlet plenum that feeds cooling air into micro-channels, for example, near a midpoint of a side of a vane. Such second inlet plenum positioning may offer the benefits of reducing the flow rate in the flow channels and reducing the flow path length for the cooling air as compared to a design utilizing a single inlet plenum. Inlet flow may enter at a midpoint of a micro-channel and split, with a portion (e.g., half) of the flow traveling upstream toward a leading edge region and the remainder of the flow traveling downstream toward the vane trailing edge. Alternatively, the second inlet plenum and micro-channels may be positioned such that all of the suction-side wall or pressure-side wall cooling air is routed in a single direction along the side wall(s).
To reiterate, in an embodiment of the presented inventions, the leading edge may be cooled by cooling air impinging the internal surface of the leading edge at some point between the stagnation point and the suction-side heating peak. A portion of the cooling air may travel through a finned array (i.e., a plurality of fins) in a leading edge flow channel toward the stagnation point, while the remaining cooling air may travel through another finned array toward the suction-side heating peak. The flow from both finned arrays may exit through outlet plenums which may be aft of the finned regions. It is also contemplated that multiple cooling air flows may be combined and exhausted through a single outlet plenum. A second inlet plenum may introduce cooling air approximately midway down the chord of the vane with flow traveling both forward and aft on one or both side walls of the vane through micro-channels, and exiting through an outlet plenum, the same or different than the outlet plenum utilized to exhaust the flow from the finned array(s). As an example, micro-channels cooling the side walls may be approximately 6 mils high in the spanwise direction, with 6 mil thick solid regions (fins) on either side, although these dimensions should not be considered limiting as micro-channels may be smaller or larger. Air distribution plenums described herein may run the entire span of a vane, as may finned regions. Alternatively, various features described herein such as inlet plenums, outlet plenums, finned regions, etc., may be disposed only in a targeted region or regions of a vane associated with high temperatures.
In another embodiment, cooling air may impinge the internal surface of the leading edge at a point directly corresponding to the stagnation point. A portion of the cooling air may travel through a finned array in a leading edge flow channel toward an outlet plenum, while the remaining cooling air may travel through another finned array toward the same, or a different, outlet plenum. In this regard, a vane design may prevent a suction-side heating peak and therefore, the stagnation point may be the sole critical heating peak requiring improved cooling.
The general cooling principles outlined herein may be combined with other cooling techniques. For example, cooling air may be exhausted through film-cooling exhaust ports after passing through a finned array in the leading edge flow channel or a micro-channel instead of routed through an outlet line in a manifold. In such an arrangement, the mass flow of the cooling air discharged into the core flow of the turbine may be reduced in comparison to prior film-cooling arrangements.
The presented inventions provide several advantages over prior art solutions including but not limited to: 1) an ability to operate at current engine operating temperatures using only internal cooling, thereby eliminating film-cooling and the associated aerodynamic penalties; 2) turbine inlet temperature capability may be increased by 350° F. or more during operational conditions, while eliminating film-cooling; and 3) vane leading edge cooling efficiency may be increased more than 20% when compared to conventional vane performance. These improvements may yield significant advances in fuel efficiency. The benefits stemming therefrom may include: 1) potential cost savings that presently correspond to several billion dollars worldwide, 2) improved efficiency directly translating into lower CO2 emissions and decreasing the impact of CO2 emissions on the environment (additionally, more complete combustion will occur at higher inlet temperatures resulting in lower emissions of SOx and NOx), 3) increased engine life allowing airlines to increase the mean time before overhaul (MTBO) which may translate into more uptime hours, and 4) increased turbine reliability allowing airlines to operate with a higher reliability factor which may translate into more on-time departures and fewer hours of out-of-service repairs.
Numerous additional features and advantages of the presented inventions will become apparent to those skilled in the art upon consideration of the embodiment descriptions provided hereinbelow.
The following description is not intended to limit the inventions to the forms disclosed herein. Consequently, variations and modifications commensurate with the following teachings, skill and knowledge of the relevant art, are within the scope of the presented inventions. The embodiments described herein are further intended to explain modes known of practicing the inventions and to enable others skilled in the art to utilize the inventions in such, or other embodiments and with various modifications required by the particular application(s) or use(s) of the presented inventions. Although the term “vane” is frequently used herein, and the features and functionality may be primarily directed to turbine vanes (i.e., stators), it should be appreciated that the internal cooling concepts described may be equally applicable to a rotating turbine blade or other gas turbine engine component and use of the term “vane” should not be interpreted as limiting the scope of this disclosure. Rather, the word vane is used because nozzle guide vanes disposed nearest the combustor are often subjected to the most extreme heat and may be prone to failure, thus making the concepts described herein pertinent. Nevertheless, the concepts described herein may be equally applicable to other engine components and the words “vane,” “stator,” “blade,” etc. may be used interchangeably and may refer to any gas turbine engine component subjected to high temperatures which require some type of cooling system.
The highest temperatures in a gas turbine engine are commonly encountered at the nozzle guide vanes 106 and high pressure turbine components 104. These high temperatures are typically a critical limiting factor in the operation of a gas turbine engine due to material limitations. That is to say, it is desired in the field to utilize metals or other materials with exceptional temperature tolerance, to improve cooling systems, or both. Many gas turbine engines must be throttled back from their full power generation capacity so as to avoid extreme temperatures which may cause material failure such as cracking or softening in the nozzle guide vanes 106 and high pressure turbine 104 components. Improved materials of construction or cooling of these critical features may allow gas turbine engines to operate at higher temperatures yielding improved performance and efficiency.
Typically, the highest temperatures experienced on the outer surface of a vane 300 occur in the vicinity of the stagnation point 301 due to the flow being brought to rest at the stagnation point 301. Another region of high temperatures may also form at a suction-side heating peak location 302 due to the flow of combustion gases, and their associated boundary layer, over the surface of the vane 300 transitioning from laminar to turbulent, thereby increasing the external convective heat transfer coefficient.
Cooling air may be circulated internally in a vane 400 to draw heat away from the outer wall 417 which may otherwise become overheated in the absence of impingement cooling and/or film-cooling. Cooling air may enter a given stratum of the vane 400 through a first inlet plenum 401 which passes longitudinally through at least a portion of the vane 400 to distribute the cooling air throughout the vane 400. The cooling air may be pressurized to promote rapid air flow for improved cooling. In the illustrated embodiment, the cooling air may exit the first inlet plenum through an impingement nozzle 419 and impinge upon a wedge 402 which bifurcates the cooling air flow with a portion flowing into a flow channel 412 which passes across the suction-side heating peak 403 and a portion flowing into a flow channel 413 which passes across the stagnation point 404. The flow channels 412, 413 are defined by the inner surface of the outer wall 417 and an outer surface of one or more inner walls 418b, 418c. The wedge 402 prevents the formation of a stagnation point on the internal side of the outer wall of the vane 400 and encourages smooth flow entry into the flow channels 412, 413. The wedge 402 may also define the impingement angle of the cooling air flow upon the outer wall to maintain cooling air flow velocity, wherein a wide wedge may translate to higher velocities through the flow channels than a narrow wedge. However, it will be appreciated that the wedge 402 is not required. As cooling air passes across the inner surface of the outer wall 417 of the vane 400, it convectively draws heat from the outer wall 417 and carries it away toward an outlet. Heated cooling air may be circulated from flow channel 412 toward outlet plenum 405 and from flow channel 413 toward outlet plenum 406.
Similarly, cooling air may also be brought into circulation through a second inlet plenum 407 and directed to cool various targeted surfaces. For example, a portion of the cooling air may be circulated through micro-channel 414 to cool the suction-side outer wall 408 before being exhausted through outlet plenum 405 and a portion of the cooling air may be circulated through micro-channel 415 to cool the pressure-side outer wall 409 before being exhausted through outlet plenum 406. A portion of the cooling air entering through inlet plenum 407 may also be directed rearward. For example, a portion of the cooling air may be circulated through micro-channel 411 to cool the suction-side outer wall 408 and a portion of the cooling air may be circulated through micro-channel 416 to cool the pressure-side outer wall 409 before being exhausted through outlet plenum 410. It should be appreciated that the described embodiment is only exemplary in nature and any combination of flow channels, micro-channels, inlet plenums, and outlet plenums may be utilized. For example, one inlet plenum and one outlet plenum may be used or three inlet plenums and two outlet plenums. The number of inlet and/or outlet plenums that may be incorporated into a design may be related to a desired engine operating temperature, a volume of cooling air needed to sufficiently cool the vane given the desired engine operating temperature, the heat capacity of the cooling air, the mass flow rate of the cooling air, and the surface area available for impingement cooling, etc. It should be appreciated that having multiple inlet plenums disposed at different locations may allow for unheated cooling air to be directed to the highest temperature areas of the vane more effectively than if only a single inlet plenum were used and the cooling air was circulated around the entirety of the inner surface of the outer wall without the introduction of additional cooling air in secondary locations.
A micro-structured surface may comprise longitudinal grooves (i.e., running in the direction of cooling air flow). It is contemplated that grooves need not run longitudinally but may provide less resistance to air flow. In either regard, the grooves may provide additional impingement surface area equal to the area of the two side walls which form each groove. For example, a given area of a wall (e.g., 1 square inch) with grooves having a depth equal to their width (w), and having a spacing (g) equal to their width, would provide approximately twice as much surface area (e.g., 2 square inches) as a smooth surface of the same area (i.e., 1 square inch). Each groove may be of any suitable height, for example, 0.1 mm or 5 mm. Similarly, spacing between grooves (g) may be modified to achieve desired cooling air flow and cooling properties.
Fins, as discussed herein, are protrusions extending from the surface of a wall. Fins may generally be any suitable shape such as round, flat, square, etc. In
The inventors have determined that L2 values smaller than the L1 dimension may increase the pressure drop observed from the inlet nozzle to the exit of a leading edge flow channel, leading to lower surface velocities and higher maximum vane material temperatures. Moreover, the pressure drop through fins is less than that associated with a similar groove design.
An impingement surface may be micro-textured in lieu of or in addition to fins, grooves, etc. Micro-texturing may comprise a chemical or mechanical process of roughening a surface to increase the available surface area.
A micro-structured surface may comprise a flow confinement shroud (not shown in
The pressure drops observed in a flow channel are the largest when the fins extend all the way to the confinement wall and the channel is the smallest. As would be expected, tests have shown that the best heat transfer occurs when the cooling air flow velocity is maximized. Thus observation of the velocity, temperature, and pressure contours indicated that it may be desirable for fins to span the entire distance from the impingement wall (e.g., inner surface of the outer wall of a vane) to the confinement wall (i.e., shroud). This design may lower the vane material temperature because all of the cooling air in the flow channel is confined and forced through the fins. However, a significant pressure drop may occur in this scenario as the flow area is reduced. It may be desirable to create high velocity flow in the hottest regions of the flow channel and sacrifice extended surface area for heat transfer in the cooler regions.
A micro-structured surface (e.g., surfaces 602, 603, 604) may provide a significantly greater amount of surface area as compared to a reference smooth surface of the same general size (e.g., surface 601). For example, a micro-finned surface may provide 5 times more surface area, 10 times more surface area, 20 times more surface area, etc. as compared to a smooth surface having the same overall height and width. As determined by simulation models, the total heat flux of an embodiment of a shrouded fin design similar to micro-finned surface with a shroud 604 was 2.19 times higher than a smooth surface impingement design such as reference smooth surface 601. This improvement may be attributable to the additional surface area provided and may translate directly into higher allowable combustor temperatures or a lesser volume of cooling air flow needed.
As previously discussed, embodiments of the presented inventions may include thin foils which may be stacked and bonded together to define a vane (or other gas turbine engine component).
The flow channels are typically very narrow. For instance, the distance between the inside surface of the outer wall and the outside surface of the inner wall in the leading edge region is between about 0.8% to about 4.5% of the chord length of the blade. In one embodiment, this distance is constant throughout the entire leading edge region. Similar or identical distances are utilized between the outer and inner walls in the sidewall micro-channels as well. These distances were chosen as a good tradeoff between pressure drop and thermal performance. The smaller the spacing, the higher the flow velocity, which produces higher heat transfer coefficients, and thus more effective cooling. However, when going to smaller wall spacings, there is an increase in the pressure drop. The converse is true for larger wall spacings.
Foils may stacked in any appropriate order to provide a desired flow channel size, for example, a stack of foils may alternate between one first foil 700 and one second foil 701. Alternately, a stack of foils may comprise three first foils 700 alternated with three second foils 701. It should be appreciated that it may be desirable to provide flow channels along the leading edge of every foil to maximize cooling in that region while micro-channels adjacent to the suction-side wall and pressure-side wall may only be provided in some foils and not in others to produce a conduction fin effect in the sidewall regions and reduce the amount of cooling air circulated to that which is needed. It should be appreciated that more than two foil configurations (e.g., first foil 700 and second foil 701) may be used to construct a vane.
As illustrated in
Half-thickness fins may be formed by any appropriate means of construction including but not limited to laser-etching, chemical etching, mechanical grinding, 3D printing, etc. In an embodiment, the etching process may use a mixture that includes hydrochloric and/or nitric acid. Of note, in an embodiment of metal foils using Hastelloy X, it has been determined by the inventors that a process for etching which provides good etch quality with feature sizes on the order of those described herein may require a more aggressive chemical etchant than is typical for other metals and may require extra passes for shorter times through the etcher.
In the case of 0.002″ foil thicknesses, half-etching of the fins may yield an approximately 0.001″ thick fin. Tabs may be used along the side wall flow channels to support the internal structures such as the internal walls and may also be half-etched (tabs may also be used in the leading edge sections for full-thickness designs to support the internal structures). Tabs may only be needed during etching, stacking, and bonding of the foils because the internal structures may become bonded to the manifolds during bonding and receive support therefrom. If the tabs were not present prior to bonding, the internal structures of the vane would likely fall out or become misaligned during assembly. It is desirable to remove these tabs after bonding but this may not always be possible and, therefore, the tabs may remain inside a completed vane in some embodiments. In such instances, the location and shape of each tab may be designed such that they present minimal impact in terms of pressure drop and flow disturbance. In alternative manufacturing techniques such as 3D printing, these tabs may not be necessary and may be omitted.
It should be noted that many of the illustrations provided in the figures depict a vane having a single cross-sectional profile. In other words, the shape of the outer perimeter of each foil may be identical. However, it should be appreciated that in many real-world implementations, nozzle guide vanes and other gas turbine engine components may be clocked or twisted such that moving in the spanwise direction, each cross-sectional profile may be unique. In this regard, a 2.4″ vane using 0.002″ foils may require up to 1200 unique foil patterns.
Alternatively to half-thickness fins, a vane may be constructed with full-thickness leading edge fins. In this design, at least two staggered finned arrays may be needed at the leading edge. However, as opposed to the half-etched version, full-thickness fins would have no spanwise space between the staggered fin layers if they were simply stacked in an A-B-A-B fashion. Therefore, a leading edge spacer foil may be needed to provide a sufficient cross-sectional area for air flow. Such a leading edge spacer foil may provide room for the cooling air to pass between layers of fins. Because of this additional type of leading edge design, as you travel in the spanwise direction, the leading edge fins may alternate in an A-C-B-C-A-C-B-C repeating pattern. In addition to the leading edge fin pattern, side wall micro-channels may be sized to be 0.006″ high in the spanwise direction. Using 0.002″ thick foils, the foils may be stacked as they were described in relation to the half-etched version with three foils in a row with side wall micro-channels, followed by three foils in a row that do not have side wall micro-channels (i.e., side wall spacer foils).
The process for stacking the foils may include using alignment pins. The tight tolerance of these pins, along with a combination of etched circular and elongated holes in each foil, may allow the features on each successively stacked foil to be precisely aligned. The stacking fixture used to hold the stack of foils may be inserted into a bonding fixture. As an example, a bonding fixture may consist of two molybdenum platens connected with molybdenum threaded rods. The bonding fixture may then be placed in a vacuum furnace and ramped to 2100° F. for several hours. Due to the thermal expansion mismatch, the foil may expand at a greater rate than the molybdenum bonding fixture initially. This may help ensure the foils are in intimate contact, enabling good diffusion bond strength. Due to a minute amount of variance in the thicknesses of the foils, and due to the presence of some amount of “crush” experienced during the bonding phase (e.g., about 1.25%), an additional amount of foil may be stacked to account for this while still hitting a target vane length.
Following bonding of the foils, the vanes may be removed from the bonded stack through wire electrical discharge machining (EDM). This may consist of an electrified wire which produces a series of rapidly recurring current discharges between the wire and the part being machined, effectively burning away a small amount of material with each spark.
In an embodiment, a plurality of inner manifolds and/or a plurality of outer manifolds may be fluidly interconnected such that a single manifold is operative to feed or collect cooling air to or from a plurality of vanes. In another embodiment, only a single manifold may be used on a vane such that an inlet line and an outlet line extend from the same manifold. It should be noted that although the terms “cooling air” and “air” are used herein to describe the fluid circulated through a turbine component for cooling, it should be appreciated that any suitable cooling fluid may be utilized. Particularly in the case of a cooling system in which the cooling fluid is not dumped into the core flow of the engine but is returned through an outlet line, alternative cooling fluids may be used in place of cooling air. In such an instance, cooling fluid may be supplied from a reservoir rather than fed from the compressor. It should also be appreciated that although shown as separate components in
Testing conducted using inventive features described herein indicated an increase in the turbine inlet temperature of a gas turbine engine of 345° F. or more above current operating conditions. The result of this improvement may be harnessed by either reducing the cooling air flow rate and keeping the turbine inlet temperature constant, thereby reducing the cycle losses from pulling air off the high compressor outlet, or by keeping the cooling air flow rate constant and increasing the turbine inlet temperature, thereby increasing the power and cycle efficiency.
The foregoing description of the presented inventions has been presented for purposes of illustration and description. Furthermore, the description is not intended to limit the inventions to the forms disclosed herein. Consequently, variations and modifications commensurate with the above teachings, and skill and knowledge of the relevant art, are within the scope of the presented inventions. The embodiments described hereinabove are further intended to explain known modes of practicing the invention and to enable others skilled in the art to utilize the inventions in such or other embodiments and with various modifications required by the particular application(s) or use(s) of the presented inventions. It is intended that the appended claims be construed to include alternative embodiments to the extent permitted by the prior art.
This application claims benefit of U.S. Provisional Patent Application No. 62/534,505 filed Jul. 19, 2017, entitled “MICROTEXTURED SURFACES FOR TURBINE BLADE IMPINGEMENT COOLING,” which is incorporated herein by reference in its entirety.
The presented inventions were made with government support under Contract #NNX12CA18C awarded by the National Aeronautics and Space Administration. The government has certain rights in the inventions.
| Number | Date | Country | |
|---|---|---|---|
| 62534505 | Jul 2017 | US |
