FOR A HYBRID AIRFOIL

FIELD OF THE INVENTION

[0001] This invention relates to airfoils, such as vanes or blades, used in a gas turbine, and more particularly, to an airfoil having at least one bumper element located within a surface gap located between an outer shell and core support structure of the airfoil, wherein the bumper element provides a thermally conductive path for conducting heat from the outer shell to the core support structure and wherein the bumper element is flexible to absorb relative movement between the outer shell and the core support structure to reduce stress levels in the airfoil..

BACKGROUND OF THE INVENTION

[0002] In various multistage turbomachines used for energy conversion, such as gas turbines, a fluid is used to produce rotational motion. Referring to Fig. 1, an axial flow gas turbine 10 includes a multi-stage compressor section 12, a combustion section 14, a multi stage turbine section 16 and an exhaust system 18 arranged along a center axis 20. Air at atmospheric pressure is drawn into the compressor section 12 generally in the direction of the flow arrows F along the axial length of the turbine 10. The intake air is progressively compressed in the compressor section 12 by rows of rotating compressor blades, thereby increasing pressure, and directed by mating compressor vanes to the combustion section 14, where it is mixed with fuel, such as natural gas, and ignited to create a combustion gas. The combustion gas, which is under greater pressure, temperature and velocity than the original intake air, is directed to the turbine section 16. The turbine section 16 includes a plurality of airfoil shaped turbine blades 22 arranged in a plurality of rows Ri, R ₂ , etc. on a shaft 24 that rotates about the axis 20. The combustion gas expands through the turbine section 16 where it is directed in a combustion flow direction F across the rows of blades 22 by associated rows of stationary vanes 24. A row of blades 22 and associated row of vanes 24 form a stage. In particular, the turbine section 16 may include four stages. As the combustion gas passes through the turbine section 16, the combustion gas causes the blades 22 and thus the shaft 24 to rotate about the axis 20, thereby extracting energy from the flow to produce mechanical work. [0003] A method for increasing the efficiency of a turbine is to increase an operating temperature of the turbine. Operation of a turbine at higher temperatures frequently results in the use of turbine components such as vanes and/or blades (i.e. airfoils) fabricated from more than one material to form a hybrid configuration. For example, an airfoil may include a high heat resistant material used in combination with an alloy support structure. In this configuration, the high heat resistant material may become stressed during operation of the gas turbine due to different rates of thermal expansion between the high heat resistant material and the alloy material. It is desirable to minimize the effects of thermal expansion between the high heat resistant material and alloy support structure.

SUMMARY OF INVENTION

[0004] An airfoil for a turbine is disclosed. The airfoil includes a core support structure having a core outer surface and a high heat resistant outer shell having leading and trailing edges and high and low pressure surfaces. The outer shell also includes a shell inner surface, wherein a portion of the core support structure is located within the outer shell and wherein the core outer surface and shell inner surface are separated by a surface gap. Further, the airfoil includes at least one bumper element located within the surface gap, wherein the at least one bumper element provides a thermally conductive path for conducting heat from the outer shell to the core support structure and wherein the at least one bumper element is flexible to absorb relative movement between the outer shell and the core support structure to reduce stress levels in the airfoil.

[0005] Those skilled in the art may apply the respective features of the present invention jointly or severally in any combination or sub-combination. BRIEF DESCRIPTION OF DRAWINGS

[0006] The teachings of the present disclosure can be readily understood by considering the following detailed description in conjunction with the accompanying drawings, in which:

[0007] Fig. 1 is a partial cross sectional view of an axial flow gas turbine.

[0008] Fig. 2 i s a cross sectional view of an exemplary airfoil.

[0009] Fig. 3 is an expanded view of section 3 of Fig. 2 depicting an expanded cross sectional view of the airfoil.

[0010] Fig. 4 depicts an alternate embodiment of a bumper element.

[0011] Fig. 5 depicts a still further embodiment of a bumper element.

[0012] Figs. 6 A and 6B illustrate computer modeling of body temperature distribution for a thermal barrier coating (TBC) layer, outer shell and support structure when using bumper elements (Fig. 6 A) and without bumper elements (Fig. 6B) and when subjected to an elevated temperature located adjacent the TBC layer.

[0013] Figs. 7 A and 7B illustrate computer modeling for Von Mises stresses and A maximum strength criteria parameter, respectively, for the TBC layer and outer shell.

[0014] To facilitate understanding, identical reference numerals have been used, where possible, to designate identical elements that are common to the figures.

DETAILED DESCRIPTION

[001 5] Although various embodiments that incorporate the teachings of the present disclosure have been shown and described in detail herein, those skilled in the art can readily devise many other varied embodiments that still incorporate these teachings. The scope of the disclosure is not limited in its application to the exemplary embodiment details of construction and the arrangement of components set forth in the description or illustrated in the drawings. The disclosure encompasses other embodiments and of being practiced or of being carried out in various ways. Also, it is to be understood that the phraseology and terminology used herein is for the purpose of description and should not be regarded as limiting. The use of "including," "comprising," or "having" and variations thereof herein is meant to encompass the items listed thereafter and equivalents thereof as well as additional items. Unless specified or limited otherwise, the terms "mounted," "connected," "supported," and "coupled" and variations thereof are used broadly and encompass direct and indirect mountings, connections, supports, and couplings. Further, "connected" and "coupled" are not restricted to physical or mechanical connections or couplings.

[0016] Referring to Fig. 2, a cross sectional view of an exemplary airfoil 30 is shown, it is understood that aspects of the present invention are applicable to any airfoil having a high heat resistant outer portion and an inner core support structure (i.e. a hybrid airfoil ) and that other configurations for a hybrid airfoil in addition to that shown in Fig. 2 and described herein may be used. The airfoil 30 may be either a vane or blade used in a gas turbine 10. The airfoil 30 may include a high heat resistant outer shell portion 32 and a core support structure 34 located substantially inside the outer shell 32 to form a hybrid airfoil configuration. In this arrangement, the outer shell 32 is subjected to a substantial portion of thermal loading whereas the core support structure 34 is subjected a substantial portion of mechanical loading. The outer shell 32 includes leading 36 and trailing 38 edge outer surfaces and a first shell outer surface having a substantially concave shaped profile to provide a high-pressure surface 40. In addition, the outer shell 32 includes a second shell outer surface having a substantially convex shaped profile to provide a low- pressure surface 42. The core support, structure 34 includes a core outer surface 44 having a shape that corresponds to a shape of a shell inner surface 46. The core outer surface 44 and shell inner surface 46 are separated by a surface gap 48. The core support structure 34 also includes an airfoil cavity 52 defined by a core inner surface 50. in an embodiment, the core support structure 34 may include stiffening elements 54 that extend across the airfoil cavity 52 and between first 50 A and second SOB portions of the core inner surface 50.

[0017] A rate of thermal expansion of the high heat resistant material used to fabricate the outer shell 32 is different than a rate of thermal expansion of the alloy material used to fabricate the core support structure 34. Therefore, portions of the airfoil 30 are subjected to relatively high stresses during operation of the gas turbine 10 due to a difference in thermal expansion rates. For example, the core support structure 34 may be fabricated from an alloy suitable for providing structural support in a turbine vane 22 or blade 24 such as ESTCO EL® alloy available from Special Metals Corporation, New Hartford, NY, US. The outer shell 32 may be fabricated from a high heat resistant material such as ceramic matrix composite (CMC) material which is typically weaker and more brittle than an alloy. The CMC material and the alloy have different coefficients of thermal expansion. Thus, undesirable stress levels may be generated in the CMC material during operation of the gas turbine 10 due to different rates of thermal expansion between the CMC material and the alloy. Further, additional stresses are imposed on the CMC material due to high pressure loads from active cooling of the airfoil 30.

[001 8] In accordance with aspects of the present invention, at least one bumper element 56 is located in the surface gap 48 between the core outer surface 44 and the shell inner surface 46. Each bumper element 56 serves to conduct heat from the lead 36, trailing 38, high 40 and low 42 pressure surfaces of the outer shell 32 to the core support structure 34 so as to reduce a difference in thermal expansion between the outer shell 32 and the core support structure 34. In Fig. 2, an end view of a plurality of bumper elements 56 is shown.

[0019] Fig, 3 is an expanded view of section 3 of Fig. 2, In an embodiment, a known thermal barrier coating (TBC) layer 68 may be formed on the lead 36, trailing 38, high 40 and low 42 pressure surfaces of the outer shell 32 or portions thereof. The TBC layer 68 serves to thermally insulate the outer shell 32 from an elevated outside temperature adjacent the TBC layer 68. Each bumper element 56 includes an upper contact surface 58 that contacts portions of the shell inner surface 46. The core outer surface 44 includes cutout portions 66 that receive a lower contact surface 60 of an associated bumper element 56. The cutout portions 66 serve to locate the bumper element 56 on the core outer surface 44 in a desired position. The bumper elements 56 may be equal ly spaced relative to each other or the spacing between bumper elements 56 may vary in order to provide desired heat transfer.

[0020] In accordance with aspects of the present invention, each bumper element 56 provides a thermally conductive path for conducting heat from the outer shell 32 to the core support structure 34 so as to reduce a difference in thermal expansion between the outer shell 32 and the core support structure 34. In addition, each bumper element 56 is separated from an adjacent bumper element 56 by a bumper gap 62, thus forming a space or path that enables convection cooling between the outer shell 32 and the core support structure 34. The surface gap 48 may be sized smaller than each bumper element 56 to form an interference fit that holds each bumper element 56 between the core outer 44 and shell inner 46 surfaces. Alternatively, each bumper element 56 may be unistructurally or integral ly formed as one piece with the core support structure 34 by using a casting process or three dimensional (3D) printing techniques. Further, each bumper element 56 is fabricated from a thermally conductive material having resilient or flexible characteristics. Thus, each bumper element 56 is able to change in size in response to relative movement between the outer shell 32 and the core support structure 34, thereby absorbing the relative movement and reducing stresses in the airfoil 30.

[0021] In an embodiment, each bumper element 56 is shaped in a spiral configuration that forms a coil. For example, each bumper element 56 may be formed by roiling a sheet of thermally conductive material such as INCONEL® alloy, copper-nickel alloy or other suitable thermally conductive alloy. Referring to Fig. 4, an alternate embodiment of a bumper element 70 is shown. In this embodiment the bumper element 70 includes a plurality of alternating upper 72 and lower 74 surfaces connected by a transition section 76 to form a wave like or substantially corrugated shape. Each upper 72 and lower 74 surface contacts the shell inner 46 and core outer 44 surfaces, respectively, to provide a thermally conductive path. Referring to Fig, 5, a still further embodiment of a bumper element 80 is shown. In this embodiment, each bumper element 80 has a substantially cylindrical shape that includes a section that is separated or split 82 to form a split cylinder configuration. Each bumper element 80 includes cylinder upper 86 and lower 88 cylinder contact surfaces that contact the shell inner 46 and the cutout portions 66, respectively, to provide a thermally conductive path. Other embodiments for a bumper element include spring elements or a wire mesh arrangement such as that available from UNIQUE WIRE WEAVING CO., INC,, Hillside, NJ, US. In accordance with aspects of the invention, combinations of different types of bumper elements, spring elements and/or wire mesh may be simultaneously used in the airfoil 30.

[0022] Figs. 6A and 6B illustrate computer modeling of body temperature distribution for the TBC layer 68, outer shell 32 and support structure 34 when using bumper elements 56 (Fig. 6A) and without bumper elements 56 (Fig. 6B) and when subjected to an elevated temperature located adjacent the TBC layer 68. In particular, the temperature of the TBC layer 68 and outer shell 32 when using bumper elements 56 (see region 78 in Fig. 6A) is substantially reduced as compared to when bumper elements 56 are not used (see region 84 in Fig. 6B). It has been found that aspects of the current invention effectively cool a 4mm thick CMC layer exposed to an outside temperature of approximately 1700 degrees C. This enables a substantial reduction in the amount of cooling air needed to cool the airfoil 30.

[0023] Figs. 7A and 7B illustrate computer modeling for Von Mises stresses and maximum strength criteria parameter (must be less than 1), respectively, for the TBC layer 68 and outer shell 32. In particular, it can be seen that the Von Mises stress levels

90 and strength criteria parameter 92 when using bumper elements 56 are both within acceptable levels.

[0024] Therefore, aspects of the current invention enable higher operating temperatures in a gas turbine, resulting in a reduced cooling requirement that, in turn, results in higher power, greater efficiency, and/or reduced emissions from the gas turbine. [0025] While particular embodiments of the present disclosure have been illustrated and described, it would be obvious to those skilled in the art that various other changes and modifications can be made without departing from the spirit and scope of the disclosure. It is therefore intended to cover in the appended claims all such changes and modifications that are within the scope of this disclosure.