FIELD OF THE INVENTION

The present invention relates generally to turbine airfoils, and more particularly to airfoils with dual wall construction for use in an airfoil assembly in a turbine engine.

BACKGROUND OF THE INVENTION

In a gas turbine engine, compressed air discharged from a compressor section is mixed with fuel and burned in a combustion section, creating combustion products comprising hot combustion gases. The combustion gases are directed through a hot gas path in a turbine section comprising a series of turbine stages typically including a row of stationary vanes followed by a row of rotating turbine blades. The turbine blades extract energy from the combustion gases and provide rotation of a turbine rotor for powering the compressor and providing output power.

The combustor in a gas turbine engine operates at high temperatures, and the airfoils of the vanes and blades typically include cooling systems to remove heat from the airfoil and to prolong the life of the vane and blade components. One cooling system includes a dual or four wall construction in which the vane or blade comprises an inner wall contained within and spaced from the outer wall. The inner and outer walls are typically rigidly coupled together at one or more locations. Near- wall cooling is achieved by circulating cooling air through the near-wall cooling chambers formed between the inner and outer walls. The outer wall is exposed to the hot combustion gases, while the inner wall may be at a much lower temperature. This temperature differential results in greater thermal expansion in the outer wall as compared to the cooler inner wall, which can create stress transferred through the joints between the inner and outer walls. This thermal stress can cause cracking in the outer wall and decrease the overall life of the airfoil.

SUMMARY OF THE INVENTION In accordance with one aspect of the invention, the present disclosure provides an airfoil comprising an airfoil main body including an outer wall, a main airfoil cavity, and a plurality of ribs; and at least one inner wall located in the main airfoil cavity. The outer wall defines a leading edge, a trailing edge, a pressure side, a suction side, a radially inner end, and a radially outer tip, with the main airfoil cavity being defined between the pressure side and the suction side. The plurality of ribs extend laterally through the main airfoil cavity between the pressure side and the suction side and extend radially between the radially inner end and the radially outer tip. The at least one inner wall extends generally parallel to an inner surface of the outer wall defined by one of the pressure side and the suction side and extends radially between the radially inner end and the radially outer tip of the outer wall. The at least one inner wall comprises an upstream edge, a downstream edge, a radially inner edge, and a radially outer edge, in which the upstream edge and the downstream edge are received in a complementary slot formed in an adjacent one of the plurality of ribs. The at least one inner wall is structurally independent of the outer wall and the plurality of ribs such that the at least one inner wall is uncoupled from and movable relative to the airfoil main body.

In accordance with some aspects, the airfoil main body may be an integral structure formed by integrally casting the outer wall and the plurality of ribs. In a particular aspect, the at least one inner wall may be cast in place within the main airfoil cavity. In accordance with other aspects, joints may be defined between the at least one inner wall and the airfoil main body at locations where the at least one inner wall is adjacent to the airfoil body, in which the joints may include a thin sheet spacer material structurally separating the at least one inner wall from the airfoil main body at the joints. In a particular aspect, the thin sheet spacer material may comprise a quartz sheet.

In accordance with another aspect, the at least one inner wall may comprise at least one inner wall section adjacent to at least one of the pressure side and the suction side of the outer wall. In some aspects, the at least one inner wall may comprise a first inner wall located toward the leading edge, in which the first inner wall may comprise at least one of a first pressure side inner wall section and a first suction side inner wall section, and a second inner wall located toward the trailing edge, in which the second inner wall may comprise at least one of a second pressure side inner wall section and a second suction side inner wall section. In a particular aspect, the plurality of ribs may comprise a first rib adjacent to the leading edge, in which a downstream surface of the first rib may comprise at least one complementary slot that receives the upstream edge of the first inner wall; a mid- chord rib adjacent to a mid-section of the airfoil, in which an upstream surface of the mid-chord rib may comprise at least one complementary slot that receives the downstream edge of the first inner wall and a downstream surface of the mid-chord rib may comprise at least one complementary slot that receives the upstream edge of the second inner wall; and a third rib adjacent to the trailing edge, in which an upstream surface of the third rib may comprise at least one complementary slot that receives the downstream edge of the second inner wall. In accordance with another aspect, an outer surface of the at least one inner wall may further comprise at least one tab, and the inner surface of the outer wall may further comprise at least one groove that receives the tab.

In accordance with a further aspect of the invention, the present disclosure provides an airfoil comprising an airfoil main body including an outer wall defining a leading edge, a trailing edge, a pressure side, a suction side, a radially inner end, and a radially outer tip; a main airfoil cavity defined between the pressure side and the suction side; and a plurality of ribs extending laterally through the main airfoil cavity between the pressure side and the suction side and extending radially between the radially inner end and the radially outer tip, in which the airfoil main body is an integral structure formed by casting the outer wall and the plurality of ribs as a single piece. The airfoil further comprises a first inner wall located toward the leading edge, in which the first inner wall comprises at least one of a first pressure side inner wall section and a first suction side inner wall section; and a second inner wall located toward the trailing edge, in which the second inner wall comprises at least one of a second pressure side inner wall section and a second suction side inner wall section. Each inner wall is spaced from the outer wall. Each inner wall also extends generally parallel to an inner surface of the outer wall defined by one of the pressure side and the suction side and extends radially between the radially inner end and the radially outer tip of the airfoil. Each inner wall is structurally independent of the outer wall and the plurality of ribs such that each inner wall is uncoupled from and movable relative to the airfoil main body.

In accordance with some aspects, each inner wall may comprise an upstream edge, a downstream edge, a radially inner edge, and a radially outer edge, with the upstream edge and the downstream edge of each inner wall being received in a complementary slot formed in an adjacent one of the plurality of ribs. In accordance with other aspects, joints may be defined between the airfoil body and at least one of the first inner wall and the second inner wall at locations where the inner walls are adjacent to the airfoil body. The joints may include a thin sheet spacer material structurally separating the at least one of the first inner wall and the second inner wall from the airfoil main body at the joints. In a particular aspect, the thin sheet spacer material may comprise a quartz sheet.

In accordance with another aspect of the invention, the present disclosure provides a method of making an airfoil for a turbine engine, in which the airfoil comprises an airfoil main body including an outer wall defining a leading edge, a trailing edge, a pressure side, a suction side, a radially inner end, and a radially outer tip, a main airfoil cavity defined between the pressure side and the suction side; and at least one inner wall located in the main airfoil cavity and extending generally parallel to an inner surface of the outer wall. The method comprises the steps of: providing a plurality of cores defining the airfoil main body and the at least one inner wall; forming a shell mold around the plurality of cores; and casting the airfoil by introducing a metal into the shell mold. The plurality of cores define at least one joint between the airfoil main body and the least one inner wall, and each joint comprises a thin sheet spacer material affixed to one of the plurality of cores defining the joint. The at least one inner wall is cast in place within the main airfoil cavity, and the thin sheet spacer material structurally separates the at least one inner wall from the airfoil main body such that the at least one inner wall is uncoupled from and movable relative to the airfoil main body.

In accordance with one aspect of the method, the airfoil main body may include a plurality of ribs extending laterally through the main airfoil cavity between the pressure side and the suction side and extending radially between the radially inner end and the radially outer tip of the outer wall. The step of casting the airfoil may include casting the plurality of ribs integrally with the outer wall. In addition, the at least one inner wall may extend radially between the radially inner end and the radially outer tip of the outer wall. In a particular aspect, the at least one inner wall may comprise an upstream edge, a downstream edge, a radially inner edge, and a radially outer edge. The upstream edge and the downstream edge of the at least one inner wall may be received in a complementary slot formed in an adjacent one of the plurality of ribs.

In accordance with an additional aspect of the method, an outer surface of the at least one inner wall may further comprise at least one tab, and the inner surface of the outer wall may further comprise at least one groove that receives the tab. In accordance with a further aspect, the method may further comprise removing the plurality of cores following the step of casting the airfoil. In a particular aspect, the step of removing plurality of cores may further comprise removing the thin sheet spacer material. In accordance with another aspect of the method, the airfoil main body may include a tip wall, and the step of casting the airfoil may include casting the tip wall integrally with the outer wall. In accordance with a further aspect of the method, the thin sheet spacer material may comprise at least one flange.

BRIEF DESCRIPTION OF THE DRAWINGS

While the specification concludes with claims particularly pointing out and distinctly claiming the present invention, it is believed that the present invention will be better understood from the following description in conjunction with the accompanying Drawing Figures, in which like reference numerals identify like elements, and wherein:

FIG. 1 is a perspective view of a turbine blade constructed in accordance with an aspect of the present invention;

FIG. 2 is a cross-sectional view taken along line 2-2 in FIG. 1 ;

FIG. 3 is an enlarged view of the section indicated by box 3 in FIG. 2;

FIG. 4 is an enlarged view similar to the section shown in FIG. 3 illustrating a portion of the manufacturing process;

FIG. 5 is a cross-sectional view taken along line 5-5 in FIG. 4; and FIG. 6 is an enlarged view of the section indicated by box 6 in FIG. 5.

DETAILED DESCRIPTION OF THE INVENTION

In the following detailed description of the preferred embodiment, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, a specific preferred embodiment in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the spirit and scope of the present invention.

The present invention provides a construction for a dual wall airfoil located within a turbine section of a gas turbine engine (not shown). As used throughout, unless otherwise noted, the terms "radially inner," "radially outer," and derivatives thereof are used with reference to a longitudinal axis of the airfoil 1 1 represented by arrow R in FIG. 1 ; the term "chordal" and derivatives thereof is used with reference to a chordal line C of an airfoil 1 1 , as depicted in FIG. 2; and the terms "axial,"

"upstream," "downstream," and derivatives thereof are used with reference to a flow of combustion gases through the hot gas path in the turbine section. Referring now to FIGS. 1 and 2, an exemplary airfoil assembly 10 constructed in accordance with a first aspect of the present invention is illustrated. The airfoil assembly 10 includes an airfoil 1 1 , a platform 14, and a root 16 that is used to conventionally secure the airfoil assembly 10 to the shaft and disc assembly of the turbine section (not shown) for supporting the airfoil assembly 10 in the gas flow path of the turbine section. The airfoil 1 1 includes an airfoil main body 12 comprising an outer wall 18 defining a leading edge 20, a trailing edge 22, a pressure side 24, a suction side 26, a radially inner end 28 adjacent to the platform 14, and a radially outer tip 30. The airfoil main body 12 extends in a radial direction R from the radially inner end 28 to the radially outer tip 30. The radially outer tip 30 comprises a tip wall 31 extending laterally between the pressure and suction sides 24, 26 (see FIG. 5). The outer wall 1 8 may further comprise a plurality of film cooling holes 62 extending through a thickness of the outer wall 18.

With reference to FIG. 2, the leading and trailing edges 20, 22 are spaced axially or chordally from each other with respect to a chordal direction C, and the pressure and suction sides 24, 26 are spaced laterally from each other to define a main airfoil cavity 32. The airfoil main body 12 may further comprise a plurality of ribs 34a-e extending laterally through the main airfoil cavity 32 between the pressure and suction sides 24, 26 and extending radially between the radially inner end 28 and the radially outer tip 30. The ribs 34a-e divide the main airfoil cavity 32 into a first leading edge chamber 36a, a second leading edge chamber 36b, a midsection chamber 36c, and one or more trailing edge chambers 36d. The

embodiment of the airfoil main body 12 shown in FIG. 2 includes five ribs 34a-e, but as will be apparent to those skilled in the art, the airfoil main body 12 may comprise varying numbers of ribs.

As shown in FIGS. 2 and 3, the main airfoil cavity 32 further comprises at least one inner wall 38 extending generally parallel to an inner surface 19 of the outer wall 18 adjacent to at least a portion of the pressure and/or suction sides 24, 26. The inner wall 38 also extends radially between the radially inner end 28 and the radially outer tip 30 (see FIG. 5). In the embodiment shown in FIG. 2, the inner wall 38 comprises a first inner wall located toward the leading edge 20 comprising a first pressure side inner wall section 38a and a first suction side inner wall section 38b. The inner wall 38 further comprises a second inner wall located toward the trailing edge 22 comprising a second pressure side inner wall section 38c and a second suction side inner wall section 38d. In some aspects of the invention, the pressure side inner wall sections 38a, 38c may comprise a single, continuous wall, and/or the suction side inner wall sections 38b, 38d may comprise a single continuous wall. In other aspects of the invention, the inner wall 38 may be located only along the pressure side 24 or only along the suction side 26.

A portion of the inner wall 38 as represented by the first suction side inner wall section 38b is illustrated in detail in FIG. 3. The first suction side inner wall section 38b comprises an upstream edge 40 located toward the leading edge 20, a downstream edge 42 located toward the trailing edge 22, a radially inner edge 44 located adjacent to the root 16, and a radially outer edge 46 (see also FIG. 5). The upstream edge 40 is received in a complementary slot 48 formed in a downstream surface of adjacent rib 34a, which is located adjacent to the leading edge 20 (also referred to herein as a first rib), and the downstream edge 42 is received in a complementary slot 48 formed in an upstream surface of adjacent rib 34b, which is located adjacent to a mid-section of the airfoil main body 12 (also referred to herein as a mid-chord rib). An outer surface 50 of the first suction side inner wall section 38b comprises at least one tab 52 that is received in a complementary groove 54 formed in the inner surface 19 of the outer wall 1 8 adjacent to the tab 52.

As seen in FIG. 2, the ribs 34a-e may comprise one or more additional complementary slots (not labeled). For example, a downstream surface of the first rib 34a may further comprise an additional slot that receives, for example, the upstream edge of the first pressure side inner wall section 38a. An upstream surface of the mid-chord rib 34b may further comprise an additional slot that receives, for example, the downstream edge of the first pressure side inner wall section 38a, and a downstream surface of the mid-chord rib 34b may further comprise one or more additional slots that receive, for example, the upstream edges of the second pressure side inner wall section 38c and the second suction side inner wall section 38d. An upstream surface of the rib 34c, which is located adjacent to the trailing edge (also referred to herein as a third rib), may further comprise one or more slots that receive, for example, the downstream edges of the second pressure side inner wall section 38c and the second suction side inner wall section 38d. In addition, the outer surface (not labeled) of one or more of the first pressure side inner wall section 38a, the second pressure side inner wall section 38c, and the second suction side inner wall section 38d may each comprise one or more tabs (not labeled) that are received in the grooves (not labeled) formed in the inner surface 19 of the outer wall 1 8 adjacent to the tabs.

As described in more detail herein, the airfoil main body 12, including the outer wall 1 8 and the plurality of ribs 34a-e, is a continuous structure that is formed by integrally casting the outer wall 18 and the plurality of ribs 34a-e as a single piece. Also as described herein, the inner wall 38 is cast in place within the main airfoil cavity 32, but the inner wall 38 is structurally independent of the airfoil main body 12. For example, with reference to FIG. 3, a joint or junction between the upstream and downstream edges 40, 42 of the first suction side inner wall section 38b and first and mid-chord ribs 34a, 34b define a gap 56 such that the first suction side inner wall section 38b is uncoupled from and movable relative to first and mid- chord ribs 34a, 34b. Likewise, a joint between the tab 52 and the groove 54 defines a gap 56 such that the first suction side inner wall section 38b is uncoupled from and movable relative to the outer wall 1 8. Joints between the radially inner and outer edges 44, 46 of the first suction side inner wall section 38b and the root 14 and the tip wall 31 , respectively, also define a gap such that the first suction side inner wall section 38b is uncoupled from and movable relative to root 14 and the tip wall 31 (see FIG. 5).

With reference to FIGS. 2 and 3, the inner wall sections 38a-d are spaced from the outer wall 1 8 such that the outer surface 50 of each inner wall section 38a- d defines, with the inner surface 19 of the adjacent portion(s) of the outer wall 18 and one or more of the plurality of ribs 34a-e, a plurality of radially extending near- wall cooling chambers 58a-d. Each near-wall cooling chamber 58a-d may be further subdivided by the presence of one or more tabs 52. For example, as seen in FIG. 3, near-wall cooling chamber 58b is divided into two compartments by the tab 52. In some aspects of the invention, one or more of the inner wall sections 38a-d may also include one or more holes 60 through a thickness of the inner wall section 38a- d such that the near-wall chambers 58a-d are in fluid communication with the main airfoil cavity 32. For example, as shown in FIG. 3, the first suction side inner wall section 38b can comprise a plurality of holes 60 through a thickness of the first suction side inner wall section 38b such that the near-wall chamber 58b is in fluid communication with the second leading edge chamber 36b. Also shown in FIG. 3 are the film cooling holes 62 that may extend through a thickness of the outer wall 18 to provide film cooling to the outer wall 18. A cooling airflow comprising, for example, a portion of the compressed air discharged from the compressor, may be provided to the airfoil cavity 32 by, for example, one or more passages (not shown) in the root 16. As seen in FIG. 3, a portion of the cooling airflow Ap may flow through the holes 60 to provide impingement cooling to the inner surface 19 of the outer wall 18. In addition to the airflow supplied from the root 16, the cooling airflow Ap flowing through the holes 60 may resupply the cooling air to the near-wall chambers 58a-d for use in film cooling. Because the outer wall 18 is exposed to the hot combustion gases, the outer wall 18 operates at a higher temperature as compared to the inner wall 38. As a result, the outer wall 1 8 experiences significantly higher thermal growth as compared to the cooler inner wall 38, particularly in the chordal direction C as seen in FIG. 2. Where the outer and inner walls 18, 38 are rigidly connected together as in many conventional four-wall airfoils, the differential thermal growth between the inner and outer walls 18, 38 can create areas of high stress. This stress is avoided in the present invention by structurally isolating the inner wall 38 from the airfoil main body 12 and allowing the outer and inner walls 18, 38 to move independently.

The airfoil assembly 10 according to the present invention may be

manufactured using a variety of suitable methods, including investment casting. In one exemplary embodiment, a plurality of wax patterns may be created that define an airfoil according to the present invention. FIGS. 4 and 5 illustrate the

manufacture of a portion of the airfoil assembly 10 using the first suction side inner wall section 38b as a representative portion of the inner wall 38. FIG. 4 illustrates an enlarged portion of the suction side 26 and the first suction side inner wall section 38b similar to the view depicted in FIG. 3. FIG. 5 is a cross-sectional front view of the suction side 26 of the airfoil main body 12 and includes a portion of the platform 14 and the root 16. In each of FIGS. 4 and 5, elements of the first and second wax patterns are provided with reference numerals matching the reference numerals of corresponding elements of the manufactured airfoil assembly 10 as shown in FIGS. 1 -3. Although specific reference is made herein to features of the airfoil assembly 10, the airfoil 1 1 , and the airfoil main body 12 depicted in FIGS. 1 -5, one of skill in the art will understand that additional features may be incorporated without departing from the spirit and scope of the present invention.

In one exemplary method, a first wax pattern defining the airfoil main body 12 and at least one second wax pattern corresponding to one or more inner wall sections 38a-d are produced using known methods such as injection of wax into a pattern mold having the desired shape. Additive manufacturing techniques such as three-dimensional (3-D) printing may also be used to achieve a wax pattern having a complex shape. While wax is frequently used in investment casting, those skilled in the art will appreciate that other suitable materials such as resins, plastics, and mixtures or composites thereof may also be used.

With reference to FIGS. 1 and 2, the airfoil main body 12 defined by the first wax pattern includes an outer wall 18 defining a leading edge 20, a trailing edge 22, a pressure side 24, a suction side 26, a radially inner end 28, and a radially outer tip 30 (see also FIGS. 4 and 5). The leading and trailing edges 20, 22 are spaced chordally from each other, and the pressure and suction sides 24, 26 are spaced laterally from each other to define the main airfoil cavity 32. The portion of the first wax pattern defining the radially outer tip 30 of the airfoil main body 12 comprises a tip wall 31 extending laterally between the pressure and suction sides 24, 26 (see FIG. 5). The airfoil main body 12 further comprises a plurality of ribs 34a-e extending laterally through the main airfoil cavity 32 between the pressure and suction sides 24, 26 and extending radially between the radially inner end 28 and the radially outer tip 30 of the airfoil main body 12 (see also FIGS. 4 and 5). The outer wall 18 and ribs 34a-e are integrally cast from a single, continuous wax pattern such that the resulting airfoil main body 12, including the radially outer tip 30, is a single piece. In some embodiments, portions of the airfoil assembly 10 may also be cast integrally along with the airfoil main body 12. For example, as shown in FIG. 5, the first wax pattern may include all or part of the platform 14 and/or the root 16 (see also FIG. 1 ).

One or more second wax patterns defining the inner wall 38 are produced using known methods. With reference to FIGS. 2, 4, and 5, the second wax pattern(s) may include, for example, a first inner wall located toward the leading edge 20 comprising a first pressure side inner wall section 38a and a first suction side inner wall section 38b and a second inner wall located toward the trailing edge 22 comprising a second pressure side inner wall section 38c and a second suction side inner wall section 38d. Each inner wall section 38a-d comprises an upstream edge 40, a downstream edge 42, a radially inner edge 44, and a radially outer edge 46. As previously described, the inner wall 38 extends generally parallel to the inner surface 19 of the outer wall 18 and extends radially between the radially inner end 28 and the radially outer tip 30. Also as previously described, the pressure side inner wall sections 38a, 38c and/or the suction side inner wall sections 38b, 38d may comprise a single, continuous wall, and the inner wall 38 may be located only along the pressure side 24 or only along the suction side 26.

Following formation of the first and second wax patterns, the exemplary method continues with formation of a wax lay-up. The second wax patterns are inserted into one or more predetermined locations within the first wax pattern that are configured to receive the second wax patterns. As previously discussed and as shown in FIGS. 2-4, the portion of the first wax pattern corresponding to one or more of the plurality of ribs 34a-e may be formed such that each rib comprises one or more slots in the upstream and/or downstream surface that receive the portions of the second wax patterns corresponding to the upstream and downstream edges of each of the inner wall sections 38a-d. For example, with reference to FIG. 4, the portion of the first wax pattern corresponding to the first rib 34a and the mid-chord rib 34b may comprise complementary slots 48 that receive the upstream and downstream edges 40, 42, respectively, of the first suction side inner wall section 38b. Also as previously discussed, the portion of the second wax patterns corresponding to the outer surface 50 of one or more of the inner wall sections 38a- d may comprise one or more tabs 52 that are received in the complementary grooves 54 formed in the portion of the first wax pattern corresponding to the inner surface 19 of the outer wall 18. For example, the second wax pattern corresponding to the first suction side inner wall section 38b may comprise at least one tab 52 on the outer surface 50 that is received in the complementary groove 54 formed in the inner surface 19 of the portion of the outer wall 18 adjacent to the tab 52.

The first and second wax patterns define one or more radially extending near- wall cooling chambers. For example, as illustrated in FIGS. 2-5, the outer surface 50 of each of the inner wall sections 38a-d defines, along with the inner surface 19 of the adjacent portion(s) of the outer wall 18 and one or more of the plurality of ribs 34a-e, a plurality of radially extending near-wall chambers 58a-d. As previously discussed, each of the near-wall cooling chambers 58a-d may be further subdivided by the presence of one or more tabs 52. Although not depicted in FIGS. 4 and 5, the first and second wax patterns may include apertures corresponding to the film cooling holes 62 extending through a thickness of the outer wall 18 and holes 60 extending through a thickness of one or more of the inner wall sections 38a-d (see FIGS. 1 and 2).

A casting mold is then fabricated from the wax lay-up using known methods. For example, a ceramic slurry may be introduced into the wax lay-up comprising the first and second wax patterns to form a plurality of ceramic cores 70 within the internal empty spaces inside the wax lay-up (self-forming ceramic cores). A shell mold 72 may then be formed by coating the outer surface of the wax lay-up with multiple coats or layers of ceramic. A portion of the shell mold 72 is depicted in FIG. 4 along the suction side 26 of the airfoil main body 12.

In another exemplary method in which wax patterns are not used, the ceramic cores 70 may comprise preformed cores that are formed using, for example, 3-D printing, to generate ceramic cores 70 that define the internal structures of the airfoil 1 1 , including the inner surface 19 of the outer wall 18, the plurality of ribs 34a-e, and the one or more inner wall sections 38a-d. The preformed ceramic cores 70 are placed in a master mold pattern that defines the outer wall 18. Wax is injected into the mold cavity to generate a wax preform of the airfoil 1 1 that forms around the preformed ceramic cores 70. A shell mold 72 is then formed as previously described by coating the outer surface of the wax preform with multiple layers of ceramic.

In both methods, each joint formed between the airfoil main body 12 and the one or more inner wall sections 38a-d is lined with a thin sheet spacer material 66 as shown in FIGS. 4 and 5. In one aspect of the present invention, the thin sheet spacer material 66 comprises a quartz sheet. In other aspects, the thin sheet spacer material 66 may comprise an ultra-high melting temperature metal such as platinum-rhodium, e.g., provided as a thin foil; a high temperature ceramic material; or any other suitable material, mixture, or composite thereof that is sufficiently thermally resistant such that the thin sheet spacer material 66 remains intact during the casting process.

As shown in FIG. 4, a piece of the thin sheet spacer material 66 is placed in the complementary slots 48 to provide separation between and isolate the outer wall 18 from the upstream and downstream edges 40, 42 of the first suction side inner wall section 38b. Similarly, a piece of the thin sheet spacer material 66 is placed in the groove 54 to provide separation between and isolate the outer wall 18 from the tab 52 of the first suction side inner wall section 38b. As seen in FIGS. 5 and 6, a piece of the thin sheet spacer material 66 is placed at the joints between the radially inner and outer edges 44, 46 of the first suction side inner wall section 38b and adjacent portions of the root 16 and the tip wall 31 , respectively, to provide separation between and isolate the radially inner and outer edges 44, 46 of the first suction side inner wall section 38b from the rest of the airfoil main body 12. In this manner, the second wax patterns are separated from the first wax pattern such that the airfoil main body 12 and the one or more inner wall sections 38a-d are cast as, and remain, two separate pieces.

The thin sheet spacer material 66 is held in place at the joints between the first and second wax patterns. As shown in FIGS. 4 and 5, the thin sheet spacer material 66 may comprise one or more flanges 68 that extend beyond the joint between the airfoil main body 12 and the first suction side inner wall section 38b. These flanges 68 affix the thin sheet spacer material 66 to one or more of the ceramic cores 70 defining the joint. For example, with reference to FIG. 4, the thin sheet spacer material 66 located at the joints between the first and mid-chord ribs 34a, 34b and the upstream and downstream edges 40, 42 of the first suction side inner wall section 38b are substantially U-shaped and comprise radially extending flanges 68a such that the thin sheet spacer material 66 bridges two of the ceramic cores 70 to hold the thin sheet spacer material 66 in place during the casting process. With reference to FIGS. 4 and 5, the thin sheet spacer material 66 located at the joint between the tab 52 and the complementary slot 54 formed in the outer wall 18 illustrates another aspect of the invention in which the thin sheet spacer material 66 comprises laterally extending flanges 68b located adjacent to the radially inner end 28 and radially outer tip 30. As shown in FIG. 5 and 6, the joints between the platform 16 and the radially inner edge 44 and between the tip wall 31 and the radially outer edge 46 each comprise thin sheet spacer material 66, which may comprise a U-shape with chordally extending flanges 68c. In another aspect of the invention as depicted in FIG. 6, the thin sheet spacer material 66 may comprise a T- shape with chordally extending flanges 68c. When self-forming cores are used, the ceramic slurry forms around the one or more flanges 68 of the thin sheet spacer material 66 such that the flanges 68 are integrated into and held in place in one or more of the ceramic cores 70. When preformed cores are used, the thin sheet spacer material 66 may be inserted into the cores during the appropriate fabrication step such that one or more of the ceramic cores 70 surround the one or more flanges 68 in a similar manner and hold the thin sheet spacer material 66 in place.

The wax is then removed from the shell mold 72, for example, by heating the shell mold 72 to liquefy the wax. The hollow cavities previously occupied by the first and second wax patterns define the airfoil main body 12 and the inner wall 38, and the thin sheet spacer material 66 remains intact in the joints between the airfoil main body 12 and the inner wall 38. The shell mold 72 may be fired one or more times, for example, to remove moisture, harden and strengthen the shell mold 72, and preheat the shell mold 72 prior to casting. The airfoil assembly 10 is then cast by introducing a molten metal into the hollow cavities in the shell mold 72. The airfoil main body 12 including the outer wall 18 and the plurality of ribs 34a-e are cast integrally to form a single piece and to provide structural integrity to the airfoil assembly 10. In particular, the radially outer tip 30, including the tip wall 31 , is cast integrally with the rest of the airfoil main body 12. Portions of one or both of the platform 14 and the root 16 may also be integrally cast with the airfoil main body 12.

The one or more inner wall sections 38a-d are cast in place within the airfoil 1 1 , but the thin sheet spacer material 66 remains intact during the step of casting such that the airfoil main body 12 and the one or more inner wall sections 38a-d are structurally independent of one another, i.e. the airfoil main body 12 and inner wall sections 38a-d are structurally separated by the thin sheet spacer material 66.

Because of the barrier formed by the thin sheet spacer material 66, each of the inner wall sections 38a-d is uncoupled from the airfoil main body 12 and is able to move independently relative to each other and to the airfoil main body 12. Thermal growth of the inner wall sections 38a-d thus can occur independently of the airfoil main body 12.

Following casting of the airfoil assembly 10, the shell mold 72 and ceramic cores 70 may be removed using known methods such as dissolving or leaching out the ceramic using chemical means such as citric acid or sodium hydroxide. The step of removing the shell mold 72 and ceramic cores 70 may damage or completely remove the thin sheet spacer material 66. For example, as best seen in FIG. 3, the gap 56 defined at the joints between the first suction side inner wall section 38b and the outer wall 1 8 and ribs 34a, 34b contains no remnants of the thin sheet spacer material 66. However, it is unnecessary to remove the thin sheet spacer material 66 following casting, and in some aspects of the present invention, some or all of the thin sheet spacer material 66 may remain in the gap 56 in the finished airfoil 1 1 .

While particular embodiments of the present invention 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 invention. It is therefore intended to cover in the appended claims all such changes and modifications that are within the scope of this invention.