This invention relates generally to improvements in turbomachines, such as turbine engines. More specifically, the invention is directed to methods and articles for decreasing the flow of gas (e.g., hot gas) into selected regions within the stator-rotor assemblies of gas turbines.
The typical design of most turbine engines is well-known in the art. They include a compressor for compressing air that is mixed with fuel. The fuel-air mixture is ignited in an attached combustor, to generate combustion gases. The hot, pressurized gases, which in modern engines can be in the range of about 1100 to 2000° C., are allowed to expand through a turbine nozzle, which directs the flow to turn an attached, high-pressure turbine. The turbine is usually coupled with a rotor shaft, to drive the compressor. The core gases then exit the high pressure turbine, providing energy downstream. The energy is in the form of additional rotational energy extracted by attached, lower pressure turbine stages, and/or in the form of thrust through an exhaust nozzle.
In the typical scenario, thermal energy produced within the combustor is converted into mechanical energy within the turbine, by impinging the hot combustion gases onto one or more bladed rotor assemblies. (Those versed in the art understand that the term “blades” is usually part of the lexicon for aviation turbines, while the term “buckets” is typically used when describing the same type of component for land-based turbines). The rotor assembly usually includes at least one row of circumferentially-spaced rotor blades. Each rotor blade includes an airfoil that includes a pressure side and a suction side. Each airfoil extends radially outward from a rotor blade platform. Each rotor blade also includes a dovetail that extends radially inward from a shank extending between the platform and the dovetail. The dovetail is used to mount the rotor blade within the rotor assembly to a rotor disk or spool.
The rotor forms part of a stator-rotor assembly. The rows of rotor blades on the rotor assembly and the rows of stator vanes on the stator assembly extend alternately across an axially oriented flowpath for “working” the combustion gases. The jets of hot combustion gas leaving the vanes of the stator element act upon the turbine blades, and cause the turbine wheel to rotate in a speed range of about 3000-15,000 rpm, depending on the type of engine. (Again, in terms of parallel terminology, the stator element, i.e., the element which remains stationary while the turbine rotates at high speed, can also be referred to in the art as the “nozzle assembly”).
As depicted in the figures described below, the opening at the interface between the stator element and the blades or buckets can allow hot core gas to exit the hot gas path and enter the wheel-space of the turbine engine. In order to limit this leakage of hot gas, the blade structure typically includes axially projecting angel wing seals. According to a typical design, the angel wings cooperate with projecting segments or “discouragers” which extend from the adjacent stator element, i.e., the nozzle. The angel wings and the discouragers overlap (or nearly overlap), but do not touch each other, thus restricting gas flow.
A gap remains at the interface between adjacent regions of the nozzle and turbine blade, e.g., between the adjacent angel wing-discourager projections, when such a seal is used. The presence of the gap is understandable, i.e., the clearance necessary at the junction of stationary and rotating components. However, the gap still provides a path which can allow hot core gas to exit the hot gas path into the wheel-space area of the turbine engine.
The leakage of hot gas by this pathway is disadvantageous for a number of reasons. First, the loss of hot gas from the working gas stream causes a resultant loss in energy available from the turbine engine. Second, ingestion of the hot gas into turbine wheel-spaces and other cavities can damage components which are not designed for extended exposure to such temperatures, such as the nozzle structure support and the rotor wheel.
Attempts have been made in the past to minimize the leakage of hot gas from the working gas stream. These attempts have sometimes involved the use of coolant air, i.e., “purge air”, as described in U.S. Pat. No. 5,224,822 (Lenehan et al). In a typical design, the air can be diverted or “bled” from the compressor, and used as high-pressure cooling air for the turbine cooling circuit. Thus, the coolant air is part of a secondary flow circuit which can be directed generally through the wheel-space cavity and other inboard regions.
In one specific example, the coolant air can be vented to the rotor/stator interface. In this manner, the coolant air can function to maintain the temperature of certain engine components under an acceptable limit. Moreover, the coolant air can serve an additional, specific function when it is directed from the wheel-space region into one of the gaps described previously. This counter-flow of coolant air into the gap provides an additional barrier to the undesirable flow of hot gas out of the gap and into the wheel-space region.
While coolant air from the secondary flow circuit is very beneficial for the reasons discussed above, there are drawbacks associated with its use as well. For example, the extraction of air from the compressor for high pressure cooling and cavity purge air consumes work from the turbine, and can be quite costly in terms of engine performance. Moreover, in some engine configurations, the compressor system may fail to provide purge air at a sufficient pressure during at least some engine power settings. Thus, hot gases may still be ingested into the wheel-space cavity.
Another technique for minimizing the leakage of hot gas from the working gas stream of a gas turbine is described in U.S. Pat. No. 6,481,959 (Morris et al). This patent describes the use of a supplemental air cooling system, to inhibit ingestion of hot gases into various circumferential regions of the turbine disc cavity, e.g., the gap and wheelspace regions. The system in Morris et al includes a number of ingestion inhibiting dynamic jet orifices, located on the underside of the trailing edges of a turbine nozzle.
While the concept described in Morris et al may be suitable in some situations, there are drawbacks associated with it as well. For example, the air cooling system may require a diversion of air from the compressor, and this can compromise engine performance, as alluded to previously. Moreover, it appears that the air jets used in the system must produce an airflow momentum greater than that of the hot gas moving into the gap and wheelspace, so as to inhibit such movement. Such a system would appear to require a complex design, especially if the amount of cooling air needs to be minimized.
In view of this discussion, it should be apparent that new, relatively simple techniques for reducing the leakage of hot gases from a hot gas flow path into undesirable regions within a turbine engine or other type of turbomachine would be welcome in the art. Moreover, reduction of the cooling and cavity purge-air flow which is typically required to reduce the hot gas leakage would itself have other important benefits. For example, higher core air flow would be possible, thereby increasing the energy available in the hot gas flow path.
Any innovations designed to accomplish these goals must still adhere to the primary design requirements for a gas turbine engine or other type of turbomachine. In general, overall engine efficiency and integrity must be maintained. Any change made to the engine, or to specific features within the engine, must not disturb or adversely affect the overall hot gas and coolant air flow fields. Moreover, the contemplated improvements should not involve manufacturing steps or changes in those steps which are time-consuming and uneconomical. Furthermore, the improvements should be adaptable to varying designs in engine construction, e.g., different types of stator-rotor assemblies.
One embodiment of the invention is directed to a stator-rotor assembly, comprising at least one circumferential endwall having a trailing edge that comprises a pattern of cavities. In some embodiments, the stator-rotor assembly comprises at least one interface region between a surface of the stator and a surface of the rotor. The surfaces are separated by a gap. The stator is a nozzle or vane that comprises inner and outer circumferential endwalls; and each endwall includes at least one leading edge and one trailing edge, relative to a hot gas flow path. A trailing edge of the inner circumferential endwall comprises a pattern of cavities that are capable of impeding the entry of hot gas into a wheelspace region that adjoins the gap between the stator and the rotor.
Another embodiment relates to a turbomachine. The turbomachine comprises at least one stator-rotor assembly, having an interface region with surfaces being separated by at least one gap, as described above. The trailing edge on at least one endwall comprises a pattern of cavities that are capable of impeding the entry of hot gas into a wheelspace region, as described above and detailed further below.
An additional embodiment is directed to a method for restricting the flow of hot gas through a gap between a stator and a rotor in a turbomachine. As described herein, the stator is a nozzle or vane that comprises inner and outer circumferential endwalls; and each endwall includes at least one leading edge and one trailing edge, relative to a gas flow path. The method comprises the step of forming a pattern of cavities on at least a portion of the trailing edge of the inner endwall of the stator component. The cavities have a shape and size sufficient to impede the entry of hot gas into a wheelspace area that adjoins the gap between the stator and rotor.
Each rotor blade, e.g., blade 22, includes an airfoil 23 mounted on a shank 25, which includes a platform 26. (Some of the other detailed features of the rotor blades are not specifically illustrated here, but can be found in various sources, e.g., U.S. Pat. No. 6,506,016 (Wang), which is incorporated herein by reference). Shank 25 includes a dovetail 27, for connection with corresponding dovetail slots formed on rotor wheel 12.
Blade or bucket 22 includes axially projecting angel wings 33, 34, 50 and 90 (sometimes called “angel wing seals”), as depicted in
It is evident from
The term “interface region” is used herein to describe the general area of restricted dimension which includes gaps 76 and 77, along with the surrounding portions of nozzle 18 and blade 22. For the purpose of general illustration, interface region 92 in
In accordance with normal engine operation, combustion gas being directed into the engine along hot gas path 38 flows aftward through stator-rotor assembly 21, continuing through other stator-rotor assemblies in the engine. (Technically, the combustion gas should be referred to as “post-combustion” at this stage. Moreover, it should be understood that the “hot gas” is often a mixture of gases. While the mixture is usually dominated by post-combustion gases, it may also include various coolant injections and coolant flow, e.g. from nozzle 18 and/or from coolant air stream 98, discussed below). As the hot gas stream enters axial gap 78, a portion of the gas stream (dashed arrow 37) may escape through upper gap 76 and flow into buffer cavity 80. (In some extreme situations which would be very unusual, the hot gas could continue to move through lower gap 77 and enter wheel-space region 82). As mentioned above, coolant air, indicated by arrow 98, is usually bled from the compressor (not shown), and directed from the inboard region of the engine (e.g., wheel-space 82) into buffer cavity 80, to counteract the leakage of hot gas. The deficiencies which sometimes are present in such a gas flowpath system were described previously.
With continuing reference to
As used herein, the term “cavity” is meant to embrace a variety of depressions, indentations, channels, grooves, dimples, pits, or any other type of discrete sinkhole. In some preferred embodiments, each cavity has a curved inner surface. As described below, the cavity may have a depth that is tapered along at least one dimension.
The cavity depth D is usually in the range of about 10% to about 80% of the endwall depth (“EWD”) shown in
In some instances, a typical range of cavity depth for a land-based turbine would be in the range of about 1 mm to about 3 mm. In the case of an aviation turbine, the range may usually be in the range of about 0.2 mm to about 1 mm. Those skilled in the art will be able to select the most appropriate cavity depth for a given situation, based on the factors mentioned above, as well as fluid flow studies, discharge coefficient tests, computational fluid dynamics predictions, and the like. A typical range for cavity width is about 1 mm to about 10 mm.
As mentioned above, the cavities may be present in a variety of shapes, as shown, for example, in
Additional, non-limiting examples of cavity shapes can be provided. For example,
The cavity 160 depicted in
The cavity 170 depicted in
In terms of cavity location, reference can be made to
Moreover, the cavities may be in contact with each other, e.g., where the edge of each cavity is in contact with an edge of an adjacent cavity. Alternatively, the cavities may be spaced from each other, depending on their shape, as well as the other factors noted herein. The degree of spacing may thus vary according to many of those same factors, including, of course, the shape of the cavity. As a non-limiting illustration with reference to
Moreover, the actual border region between cavities need not be a relatively flat ridge, i.e., resulting in a very discrete border between each cavity. Reference is made to
While many of the primary embodiments are directed to the use of cavities for endwall sections on stationary portions of the stator-rotor assembly, the cavities may also be employed on some of the rotating components. Various types of rotors may be modified in this manner, e.g., unshrouded rotors; and those that have an attached shroud. In each instance, the cavities, as described previously, can be very useful for modifying the inner endwalls of such components.
Another embodiment of the present invention is directed to a turbomachine, which includes at least one stator-rotor assembly, such as those described above. Gas turbine engines (e.g., turbojets, turboprops, land-based power generating turbines, and marine propulsion turbine engines), represent examples of a turbomachine. Other types are known in the art as well. Non-limiting examples include a wide variety of pumps and compressors, which also happen to incorporate a stator-rotor assembly through which fluids (gas or liquid) flow. In many of these other turbomachine designs, new techniques for reducing the leakage of fluid from a flow path into other regions of the machine would be of considerable interest. Thus, the stator-rotor assemblies in any of these turbomachines could include patterns of cavities as described in this disclosure.
Still another embodiment of this invention is directed to a method for restricting the flow of gas (e.g., hot gas) through a gap between a stator and rotor in a turbomachine. The method includes the step of forming a pattern of cavities on at least a portion of the trailing edge of the inner endwall of the stator component. (The stator can be either a nozzle or vane, depending on the intended use of the stator-rotor assembly). The cavities have a shape and size sufficient to impede the entry of hot gas into a wheelspace area that adjoins the gap between the stator and rotor. (Those familiar with various types of turbomachines understand that a large number of stator-rotor assemblies are typically present, and each may include the modified endwalls described herein). The method can also be used to form cavities on selected regions of a rotating component in the assembly.
The cavities can be formed by a variety of methods. Non-limiting examples include machining methods, such as various milling techniques. Other machining processes which are possible include electro-discharge machining (EDM) and electro-chemical machining (ECM). In some cases, the cavities could be formed during casting of the particular component, e.g., the investment-casting of a turbine rotor or nozzle. As one example, an investment mold surface could be provided with a selected pattern of positive features, e.g., “mounds”, domes, pyramids, pins, or any other type of protrusions or turbulation. (Some of the methods for providing these features to various surfaces are described in U.S. patent application Ser. No. 10/841,366 (R. Bunker et al; issued as U.S. Pat. No. 7,302,990), which is incorporated herein by reference). The shape of the positive features would be determined by the desired shape of the cavities, which would be inverse to the positive feature. Thus, after removal of the mold, the part would include the selected pattern of cavities. Those skilled in the art will be able to readily determine the most appropriate technique (or combination of techniques) for forming the cavities on a given surface.
Computer models were generated to simulate the interaction between a hot gas flow path and a coolant purge flow. The models were based on the wheelspace cavity region of a stator-rotor assembly, for upstream vanes. (This region is similar to the general region in
The hot gas flow path was traced, using a passive scalar. The passive scalar simulated a tracing gas seeded at the flow path inlet. A scalar concentration “C=1” represented 100% hot gas; while “C=0” represented 100% coolant purge flow.
With continuing reference to
This invention has been described by way of specific embodiments and examples. However, it should be understood that various modifications, adaptations, and alternatives may occur to one skilled in the art, without departing from the spirit and scope of the claimed inventive concept. All of the patents, articles, and texts mentioned above are incorporated herein by reference.
This invention was made with Government support under contract number DE-FC26-05NT42643, awarded by DOE. The Government has certain rights in the invention.