Patent Application
20210215054
The present disclosure generally relates to the design and manufacture of components for gas turbine engines, particularly to turbine nozzles. More specifically, the present disclosure relates to compliant joint designs for turbine nozzles and additive manufacturing processes for the same.
A gas turbine engine includes a compressor, a combustor, and a turbine. The compressor provides compressed air to the combustor. The combustor mixes the compressed air with fuel, ignites the mixture, and provides combustion gases to the turbine. The turbine extracts energy from the combustion gases. The turbine includes one or more stages with each stage having an annular turbine nozzle and a plurality of rotor blades. The turbine nozzle channels the combustion gases to the rotor blades and the rotor blades extract energy from the combustion gases. The turbine nozzle includes a plurality of circumferentially spaced stator vanes (airfoils) positioned between and attached to radially inner and outer bands (end-walls). The circumferentially spaced vanes define converging channels there between through which the combustion gases are turned and accelerated toward the rotor blades.
The vanes of the turbine nozzle are subject to transient thermal cycling. Turbine vanes may sustain damage due to cracking from low-cycle fatigue (LCF) and thermo-mechanical fatigue (TMF). As the vanes heat up, they expand. LCF and TMF occur when stresses develop from the differential expansion rates of the airfoils and end-walls. Thick-to-thin wall thickness transitions, which are encountered on some turbine engine designs, may exacerbate LCF and TMF issues.
One prior art approach to mitigate LCF and TMF cracking is to decouple the airfoils from adjacent end-walls. However, this is difficult because airfoil aero loading requires a connection to the end-walls to transfer the loads. Two such exemplary prior art turbine vane constructions are: (1) designs where airfoils are attached to full end-wall rings, and (2) designs where one or more airfoils are attached to segmented end-walls that are then assembled into a full ring. The former (1) designs may have issues with LCF and TMF cracking because they lack any design features that reduce such failure mechanisms. In contrast, the latter (2) are less prone to LCF and TMF cracking but may have leakage between the segments, which may hurt specific fuel consumption (SFC) and may contribute to increased pattern factor at the combustor exit due to the allocation of cooling air that could be used for combustor cooling. Thus, the prior art designs that include an end-wall connection force a trade-off between component life and SFC.
Hence, there is a need for improved turbine nozzle designs that satisfy load-transfer requirements, yet that do not incur a penalty in either component life or SFC due to their end-wall configuration. It would additionally be desirable if such components could be manufactured using modern, rapid fabrication techniques, such as additive manufacturing. Furthermore, other desirable features and characteristics of the manufacturing methods disclosed herein will become apparent from the subsequent detailed description and the appended claims, taken in conjunction with the accompanying drawings and the preceding background.
This summary is provided to describe select concepts in a simplified form that are further described in the detailed description. This summary is not intended to identify key or essential features of the claimed subject matter, nor is it intended to be used as an aid in determining the scope of the claimed subject matter.
Disclosed herein, in one exemplary embodiment, is a turbine nozzle formed of a superalloy, and including: an annular end-wall including a pocket, the pocket defining an inner surface within the annular end-wall; a vane, the vane including an airfoil portion and a boss portion, the vane extending from the pocket such that the boss portion is enclosed within the pocket and the airfoil portion extends through the annular end-wall; and a seal within the pocket, the seal including one or more protrusions extending from the inner surface of the pocket and abutting the vane at one or both of the boss portion and the airfoil portion. Further disclosed herein are additive manufacturing methods for making such a turbine nozzle, as well as gas turbine engines that include such a turbine nozzle.
The present invention will hereinafter be described in conjunction with the following drawing figures, wherein like numerals denote like elements, and wherein:
The following detailed description is merely exemplary in nature and is not intended to limit the invention or the application and uses of the invention. As used herein, the word “exemplary” means “serving as an example, instance, or illustration.” Thus, any embodiment described herein as “exemplary” is not necessarily to be construed as preferred or advantageous over other embodiments. All of the embodiments described herein are exemplary embodiments provided to enable persons skilled in the art to make or use the invention and not to limit the scope of the invention which is defined by the claims. Furthermore, there is no intention to be bound by any expressed or implied theory presented in the preceding technical field, background, brief summary, or the following detailed description.
Unless specifically stated or obvious from context, as used herein, the term “about” is understood as within a range of normal tolerance in the art, for example within 2 standard deviations of the mean. “About” can be understood as within 10%, 5%, 1%, or 0.5% of the stated value. Unless otherwise clear from the context, all numerical values provided herein are modified by the term “about.”
Before proceeding with the detailed description, it is to be appreciated that the described embodiments are not limited to use in conjunction with a particular type of turbine engine. Thus, although the present embodiments are, for convenience of explanation, depicted and described as being implemented in a multi-spool turbofan gas turbine jet engine, it will be appreciated that it can be implemented in various other types of turbine engines, and in various other systems and environments. Moreover, although the embodiments of the inventive subject matter are described as being implemented into a turbine section of the engine, it will be appreciated that the embodiments of the inventive subject matter may alternatively be used in any other section of the engine that may benefit from the inclusion of compliant joint configurations as described herein.
In this regard,
The compressor section 104 includes an intermediate-pressure compressor 120 and a high-pressure compressor 122. The intermediate-pressure compressor 120 raises the pressure of the air directed into it from the fan 112, directing the compressed air into the high-pressure compressor 122. The high-pressure compressor 122 compresses the air still further, directing the high-pressure air into the combustion section 106. In the combustion section 106, which includes an annular combustor 124, the high-pressure air is mixed with fuel and combusted. The combusted air is then directed into the turbine section 108.
The turbine section 108 includes a high-pressure turbine 126, an intermediate-pressure turbine 128, and a low-pressure turbine 130 disposed in axial flow series. The combusted air from the combustion section 106 expands through the turbines 126, 128, 130 causing each to rotate. The air is then exhausted through a propulsion nozzle 132 disposed in the exhaust section 110, providing additional forward thrust. As each turbine 126, 128, 130 rotates, each drives equipment in the engine 100 via concentrically disposed shafts or spools. Specifically, the high-pressure turbine 126 drives the high-pressure compressor 122 via a high-pressure shaft 134, the intermediate-pressure turbine 128 drives the intermediate-pressure compressor 120 via an intermediate-pressure shaft 136, and the low-pressure turbine 130 drives the fan 112 via a low-pressure shaft 138.
The high-pressure turbine (HPT) module 126 is depicted in
The HPT nozzle 231 may include two end-walls, a nozzle outer end-wall 221 and a nozzle inner end-wall 222, as better seen in
The turbine nozzle 231 illustrated in
In particular, turning now to
As such, one structural feature of the present nozzle slip joint is that the boss portion 321 is provided at a base of the airfoil portion 323 of vane 223, and further that the cavity or pocket 315 in the end-wall 222 fits the boss portion 321. The boss portion 321 serves to capture the airfoil/vane 223 so it cannot be separated from the end-wall 222. (It should also be noted that these same features may be provided for outer radial end-wall 221, except everything in a reverse radial orientation.) This structural feature is desirable to prevent a portion of the vane 223 from being liberated and sending debris into downstream rotating components in case the vane 223 oxidizes or cracks completely through the entire midspan. If this failure mechanism is not a concern for a certain vane design, an alternate embodiment of the present disclosure could omit the boss portion 321 and simply have the airfoil portion 323 extended into the cavity or pocket 315 in the end-wall 222.
Furthermore, the present disclosure utilizes the ability of AM to produce very thin gaps between adjacent solid bodies which enables a sealed and compliant joint between two pieces. In particular, the airfoil opening 335 at the outer portion 313 includes a sealing feature 304, which is embodied in the non-limiting example of
As further illustrated in
Still further with regard to
As initially noted above, manufacturing of the above-described turbine nozzle designs is adapted for use in additive manufacturing processes to form net or near-net shaped components, namely nozzles. As such, in accordance with an exemplary embodiment,
In step 402 of the method 400, the component is formed according to the model of step 401. In one exemplary embodiment, a portion of the component is formed using a rapid prototyping or additive layer manufacturing process. In other embodiments, the entire component is formed using a rapid prototyping or additive layer manufacturing process.
Some examples of additive layer manufacturing processes include: direct metal laser sintering (DMLS), in which a laser is used to sinter a powder media in precisely controlled locations; laser wire deposition in which a wire feedstock is melted by a laser and then deposited and solidified in precise locations to build the product; electron beam melting; laser engineered net shaping; and selective laser melting. In general, powder bed additive manufacturing techniques provide flexibility in free-form fabrication without geometric constraints, fast material processing time, and innovative joining techniques. In one particular exemplary embodiment, DMLS is used to produce the nozzle in step 402. DMLS is a commercially available laser-based rapid prototyping and tooling process by which complex parts may be directly produced by precision sintering and solidification of metal powder into successive layers of larger structures, each layer corresponding to a cross-sectional layer of the 3D component.
Prior to a discussion of the subsequent method steps of
During operation, a base block 440 may be installed on the fabrication support 414. The fabrication support 414 is lowered and the delivery support 434 is raised. The roller or wiper 436 scrapes or otherwise pushes a portion of the build material 470 from the delivery device 430 to form the working plane 416 in the fabrication device 410. The laser 460 emits a laser beam 462, which is directed by the scanner 420 onto the build material 470 in the working plane 416 to selectively fuse the build material 470 into a cross-sectional layer of the article 450 according to the design. More specifically, the speed, position, and other operating parameters of the laser beam 462 are controlled to selectively fuse the powder of the build material 470 into larger structures by rapidly melting the powder particles that may melt or diffuse into the solid structure below, and subsequently, cool and re-solidify. As such, based on the control of the laser beam 462, each layer of build material 470 may include un-fused and fused build material 470 that respectively corresponds to the cross-sectional passages and walls that form the article 450. In general, the laser beam 462 is relatively low power, but with a high energy density, to selectively fuse the individual layer of build material 470. As an example, the laser beam 462 may have a power of approximately 50 to 500 Watts, although any suitable power may be provided.
Upon completion of a respective layer, the fabrication support 414 is lowered and the delivery support 434 is raised. Typically, the fabrication support 414, and thus the article 450, does not move in a horizontal plane during this step. The roller or wiper 436 again pushes a portion of the build material 470 from the delivery device 430 to form an additional layer of build material 470 on the working plane 416 of the fabrication device 410. The laser beam 462 is movably supported relative to the article 450 and is again controlled to selectively form another cross-sectional layer. As such, the article 450 is positioned in a bed of build material 470 as the successive layers are formed such that the un-fused and fused material supports subsequent layers. This process is continued according to the modeled design as successive cross-sectional layers are formed into the completed desired portion, e.g., the nozzle of step 402. It may also be noted that, in one embodiment of performing build step 402, the build direction may be preferentially in the angle/orientation alpha as shown in
Returning to
Accordingly, the present disclosure has provided various embodiments of new turbine nozzles utilizing recent advances in additive manufacturing to reduce mechanical stresses in turbine vane airfoil-to-end-wall joints. In addition, the disclosure enables improved sealing since full ring designs may be employed as opposed segmented vane designs. Embodiments of the present disclosure are therefore expected to reduce LCF and TMF cracking over the prior art and increase resulting engine service intervals without incurring penalties on SFC.
While at least one exemplary embodiment has been presented in the foregoing detailed description of the invention, it should be appreciated that a vast number of variations exist. It should also be appreciated that the exemplary embodiment or exemplary embodiments are only examples, and are not intended to limit the scope, applicability, or configuration of the invention in any way. Rather, the foregoing detailed description will provide those skilled in the art with a convenient road map for implementing an exemplary embodiment of the invention. It being understood that various changes may be made in the function and arrangement of elements described in an exemplary embodiment without departing from the scope of the invention as set forth in the appended claims.
In this document, relational terms such as first and second, and the like may be used solely to distinguish one entity or action from another entity or action without necessarily requiring or implying any actual such relationship or order between such entities or actions. Numerical ordinals such as “first,” “second,” “third,” etc. simply denote different singles of a plurality and do not imply any order or sequence unless specifically defined by the claim language. The sequence of the text in any of the claims does not imply that process steps must be performed in a temporal or logical order according to such sequence unless it is specifically defined by the language of the claim. The process steps may be interchanged in any order without departing from the scope of the invention as long as such an interchange does not contradict the claim language and is not logically nonsensical.
