WO2018031032A1

WO2018031032A1 - Blade for gas turbine engine

Info

Publication number: WO2018031032A1
Application number: PCT/US2016/046778
Authority: WO
Inventors: Gilles Carrier
Original assignee: Siemens Aktiengesellschaft
Priority date: 2016-08-12
Filing date: 2016-08-12
Publication date: 2018-02-15

Abstract

Airfoil outer wall thickness of a gas turbine engine blade is initially increased in a zone that is proximate the trailing edge and blade hub greater than comparable greatest wall thickness anywhere else along the trailing edge from outboard that zone all the way to the blade tip. The initially increased thickness zone (50) includes a transition zone (Lx) that bridges the respective airfoil outer wall thicknesses (50) proximate the hub end (42) and tip (44) of the blade (20). The initially increased outer wall thickness is then reduced to a final decreased outer wall thickness post casting. Some embodiments also incorporate pedestals (52) with compound curve fillets in the increased wall thickness zone.

Description

BLADE FOR GAS TURBINE ENGINE

BACKGROUND 1. Field

[0001] The invention relates to gas turbine engine rotating blades. More particularly, the invention relates to gas turbine engine blades having an increased airfoil outer wall thickness zone proximate the hub and trailing edge (TE), which reduces blade-cracking propensity and enhances blade service life.

2. Description of the Related Art

[0002] Repetitive or cyclic loading of gas turbine engine blades during service operation may induce metal fatigue cracks in the blade substrate. Commonly recognized fatigue mechanisms are thermo-mechanical fatigue (TMF), high-cycle fatigue (HCF) and low-cycle fatigue (LCF). TMF thermal strain is attributable to blade thermal expansion and contraction experienced during large temperature changes, such as induced during engine powering on or off or changes in engine load. TMF mechanical strains are associated with centrifugal loads during engine speed changes or output load, which are cumulative with the thermally induced strains. HCF is attributed to low-amplitude, high-frequency strains induced by blade flexure during engine operation, which can induce crack propagation during ongoing engine operation. LCF is characterized by high-amplitude, low-frequency plastic strains in regions of stress concentration.

[0003] One observed blade zone that is susceptible to LCF/TMF cracking is often in a region proximate the blade airfoil trailing edge and blade hub platform, and associated pedestals that are formed within the airfoil interior, which bridge opposed interior wall surfaces of cooled blades thus forming a trailing edge ejection slot. A crack initiating at one location within this region has propensity to grow or propagate as the blade flexes cyclically during engine shaft rotation under HCF operational conditions. Conventional wisdom for blade design is to minimize turbine engine blade airfoil outer wall thickness, in order to minimize rotating mass and thermal mass TMF influences. Under such conventional wisdom, thin airfoil wall thickness also reduces airfoil cross section, which is thought to enhance aerodynamic efficiency. Past proposed solutions to reduce crack propensity in the region proximate the blade airfoil trailing edge and blade hub platform has been to incorporate a constant or compound curve hub fillet around all or part of the airfoil wall and blade platform junction, while maintaining a relatively constant airfoil wall thickness along the entire trailing edge from hub to blade tip.

SUMMARY

[0004] In order to reduce rotating mass-induced cyclic fatigue and increase aerodynamic efficiency, an initial local thickening of the airfoil outer or side wall zone in the region proximate the TE and hub actually reduces, rather than increases, peak stress at previously observed crack locations around the hub to airfoil wall TE region. Furthermore, increasing local thickening of the airfoil wall zone along the TE from the hub to approximately eight to ten percent (8-10%) of the airfoil stand length, which in some embodiments encompass typically the first five to eight TE pedestals, reduces likelihood of cracks at pedestal/side wall junction regions. Local thickening of the TE reduces the peak stress at the pedestal outer or side wall location and crack formation, which in turn enhances component service life.

[0005] In cast turbine blades, the TE airfoil wall thickening zone not only reduces stress, but also enhances the casting alloy grain structure in a way to improve creep ductility, by changing the relative rates of solidification of the airfoil TE and the adjacent blade platform mass. The thickened TE zone reduces the solidification rates in a way that enhances alloy grain structure and ductility, which beneficially retards crack initiation and propagation rate exactly where it is needed. By providing an additional step of decreasing the thickness of the airfoil outer or side wall zone post solidification stage, an increase in aerodynamic efficiency of the turbine engine blade while maintaining the enhanced alloy grain structure and ductility may occur.

[0006] Exemplary gas turbine engine blade embodiments described herein initially increase airfoil wall thickness in a zone that is proximate the trailing edge (TE) and hub greater than the comparable greatest wall thickness anywhere else along the trailing edge from outboard that zone all the way to the blade tip. This zone that is proximate the trailing edge (TE) and hub is then decreased wherein the zone outer wall thickness is approximately the same as the total airfoil span most proximate to the hub. Some embodiments also incorporate pedestals with compound curve fillets in the wall thickness zone. [0007] In one aspect of the present invention, a turbine engine blade, including a hub with a blade platform and an elongated airfoil portion includes an outer wall, also called a side wall, delimiting a pressure side, a suction side, a leading edge, and a trailing edge on its exterior surface. The airfoil outer wall delimits an airfoil interior on its interior surface. Airfoil outer wall thickness is established between the respective interior and exterior surfaces. A proximal end of the outer wall is coupled to the blade platform from the leading edge to the trailing edge. Correspondingly, a distal end of the outer wall defines a blade tip. The airfoil defines a span dimension (also referred to as a stand length, or height) between its proximal and distal ends. In a turbine blade casting, a reduction in material of the mold allows for an integrally cast initial increased airfoil outer wall thickness, adding extra material thickness to the blade region where the exterior surface of the hub outer wall and the blade platform converge. In the initial increased thickness zone, which has a zone proximal end adjoining the blade platform and a zone distal end, the zone's proximal end outer wall thickness, excluding adjoining hub fillet thickness, is initially greater along the trailing edge for approximately eight to ten percent (8-10%) of total airfoil span than the comparable greatest wall thickness anywhere else along the trailing edge from outboard that zone all the way to the distal tip. The corresponding outer wall thickness in the increased thickness zone transitions from the thicker region proximate the hub to that of the outboard thinner region that is proximate the tip. Once a solidification stage of the casting process is complete, the additional material section from the airfoil outer wall thickness is machined off for the final decreased thickness zone that is approximately the same as the total airfoil span most proximate the hub.

[0008] These and other features, aspects and advantages of the present invention will become better understood with reference to the following drawings, description and claims.

BRIEF DESCRIPTION OF THE DRAWINGS [0009] The invention is shown in more detail by help of figures. The figures show preferred configurations and do not limit the scope of the invention.

[0010] FIG. 1 is an elevational view of a suction (convex profile) side of an turbine engine blade, in accordance with an exemplary embodiment; [0011] FIG. 2 is a detailed perspective view of the airfoil suction side trailing edge (TE), proximate the hub, of the blade of FIG. 1, and further showing the outer wall thickness zone profile and the hub fillet joining the airfoil side wall and the blade platform;

[0012] FIG. 3 is plan cross sectional view through a TE cooling slot between two pedestals, taken along 3-3 of FIG. 2;

[0013] FIG 4 is a plan cross section through a cooling slot in the thickness zone, which is similar to the perspective view of FIG. 3 showing airfoil outer wall or side wall thickness T_H of the zone proximate the TE and hub;

[0014] FIG 5 is a plan cross section through a cooling slot in the thickness zone, which is similar to the perspective view of FIG. 3, showing relative differences in increased airfoil outer wall or side wall thickness of the zone proximate the TE and hub, compared to corresponding wall thickness in the TE outboard of the zone and closer to the blade tip, represented by the phantom profile lines;

[0015] FIG 6 is a plan cross section through a cooling slot in the thickness zone, which is similar to the perspective view of FIG. 5, showing relative differences in increased airfoil outer wall or side wall thickness of the zone proximate the TE and hub with a possible contour for cast solidification rate reducer, compared to corresponding wall thickness in the TE outboard of the zone and closer to the blade tip, represented by the phantom profile lines; [0016] FIG 7 is a plan cross section through a cooling slot in the thickness zone, which is similar to the perspective view of FIG. 5, the increased airfoil outer wall or side wall thickness of the zone proximate the TE and hub removed through machining; and

[0017] FIG 8 is a flow chart illustrating an exemplary disclosed embodiment.

[0018] To facilitate understanding, identical reference numerals have been used, where possible, to designate identical elements that are common to the figures. Any reference designation "XX/YY" indicated that the associated lead line is directed to both of the elements XX and YY. The figures are not drawn to scale.

DETAILED DESCRIPTION

[0019] Exemplary embodiments of the invention are utilized in gas turbine engine rotating blades. More particularly, such blades having an initial increased airfoil outer wall or side wall thickness zone proximate the hub and trailing edge (TE), which reduces blade cracking propensity and enhances blade service life. Airfoil TE wall thickness in this zone proximate the hub may be, for example, forty to sixty percent (40-60%) greater than the comparable greatest wall thickness anywhere else along the trailing edge from outboard that zone all the way to the blade tip. In many embodiments, the TE outer wall thickness proximate the blade tip and outboard or above the zone remains constant or tapers to reduced thickness along the span or stand length to the tip. The initial increased thickness zone generally comprises eight to ten percent (8-10%) of the total blade stand height. In some embodiments, the corresponding outer wall thickness in the increased thickness zone transitions from the thicker region proximate the hub to that of the outboard thinner region that is proximate the tip for an additional five to seven percent (5-7%) of the total blade span or stand height. In some embodiments, the increased thickness zone incorporates the first five to eight TE pedestals. In some embodiments, such pedestals also incorporate compound curve fillets in the increased outer wall thickness zone or in any other desired zone. The initial increased outer wall or side wall thickness zone reduces blade cracking propensity and enhances service life. The outer wall or side wall thickness zone is then decreased to increase aerodynamic performance.

[0020] In cast superalloy blades embodiments, the initial thicker outer wall zone modification not only reduces the average stress, due to the larger bearing area, but also creates simultaneously a more desirable grain structure at this location by changing the liquid metal solidification rate by increasing thermal mass and slowing down the solidification rate at this location. The thicker casting wall in the hub/TE zone advantageously also enhances grain size. Large grains reduce the number of inter granular zones that might otherwise be susceptible to crack formation. The resulting TE wall thickness zone exhibits better creep ductility properties that further retard concentrated creep stress induced crack initiation and propagation. The TE may be made at an amount of thickness that is increased from the rest of the blade outer wall such as two times, three times the typical thickness or the like. After the blade is cast and solidified, the additional metal that is present with an initial increase in thickness of the outer wall zone may then be removed by machining, leaving a thin trailing edge. Blade gaspath surfaces are traditionally left untouched, but with the machining off of the additional material, a large solidification rate reducer can be introduced without sacrificing the aerodynamic performance of the resulting blade. [0021] Some embodiments of the invention incorporate larger pedestal compound fillets, which additionally reduce stress concentration in the lower TE zone proximate the hub, and raise the fatigue life of the blade. In some embodiments, the pedestal fillet incorporates a compound fillet, in order to maintain sufficient interior cavity volume in the blade airfoil for delivery of blade coolant to the trailing edge, while still minimizing the stress concentration.

[0022] FIGs. 1-3 show an exemplary gas turbine engine rotating blade 20 embodiment. The blade 20 has a hub portion 22, a blade platform 24, and an airfoil portion 26, which is coupled to the hub portion 22 along the blade platform 24. The airfoil 26 has a leading edge (LE) 28 and a trailing edge (TE) 30. The airfoil outer profile includes a suction (convex) side 32, and a pressure (concave) side 34. The airfoil has an outer wall 36, which is also sometimes referred to as a side wall, which delimits an outer wall exterior surface 38 an outer wall interior surface 40, an airfoil hub end 42, and an airfoil tip 44. Airfoil span or stand dimension, L, which is also referred to as airfoil height or airfoil length dimension, is delimited by the airfoil outer wall 36 from its hub end 42 to the end of its blade tip 44. In certain embodiments, the blade 20 is cooled, and in other embodiments, the blade 20 is not cooled. [0023] As shown in FIG 2, the blade 20 may have a hub fillet 46 circumscribing and joined to the airfoil outer wall exterior surface 38 at the proximal or hub end 42, which is also joined to the blade platform 24. In a turbine blade casting, the hub fillet 46 may be integrally cast with the airfoil outer wall exterior surface 38 and the blade platform 24, adding extra material thickness to the blade region where the exterior surface of the hub outer wall and the blade platform 24 converge initially. However, as shown in the dotted lines 38 of FIG. 2, the airfoil outer wall exterior surface 38 is defined as running continuously from the blade platform 24 to the blade tip 44, with the hub fillet 46 being considered as additional material thickness outboard of that exterior surface. While cast superalloy blade embodiments have been discussed so far, blade embodiments herein include blades cut from a homogeneous billet or forgings, fabricated from joined subcomponents, or fabricated by sequential layer additive manufacturing techniques, such as 3-D printing.

[0024] Referring to FIGs. 3-7, the blade 20 defines a trailing edge cooling slot 48, which is formed between opposed interior walls or surfaces 40 of the airfoil outer wall 36. The TE cooling slot 48 is in communication with passages or cavities within the airfoil interior delimited by the outer wall interior surfaces 40, for passage of cooling fluid, such as water, steam or compressed air, out of the blade TE 30. In certain embodiments, a plurality of elongated pedestals 52 may span the airfoil interior, and may be oriented along the airfoil stand length L between the proximal (hub) 42 and distal (blade tip) ends 44. Other pedestals may be oriented along the TE 30 to the blade tip 44. Depending upon a particular blade design, additional airfoil interior wall spanning pedestals are oriented within other areas of the airfoil interior. Cooling passages 62 may be formed between opposed pedestals 52 in the TE cooling slot 48, so that cooling fluid can communicate with and cool the trailing edge 30.

[0025] As previously noted, the airfoil 26 has an initial increased airfoil side wall or outer wall 36 thickness zone 50, excluding any adjoining hub fillet 46 thickness, which is also proximate the trailing edge 30. As shown in FIG. 1 , the initial increased airfoil outer wall thickness zone 50 has a proximal end, which initiates at the airfoil proximal or hub end 42 and runs upwardly along the blade stand for approximately eight to ten percent (8-10%) of the total airfoil span, or stand length dimension, which is denoted by L_H. The zone L_H comprises the greatest initial increased outer wall thickness within the zone 50. Also included in the initial increased outer wall thickness zone 50, above or outboard the span or stand length dimension, L_¾ may be a transition zone span or stand length dimension, denoted by Lx, where the corresponding outer or side wall thickness decreases incrementally until merger with the outer wall's distal trailing edge zone, to the blade tip end 44, which is denoted by L_T. The initial increased thickness zone 50 distal end terminates at the outboard-most end of the transition zone L_x. The distal tip trailing edge portion, outboard of the initial increased wall thickness zone 50, which is denoted by L_T, constitutes between approximately eighty-three to eighty-seven percent (83-87%) of total airfoil span or stand length dimension, L in certain embodiments. In an exemplary embodiment, the total blade TE 30 span or stand length L is 258 mm. The initial increased outer wall thickness zone 50 span or stand length L incorporates the full increased wall thickness portion LH span of 22mm as well as the transition zone span Lx of 18mm. The remaining trailing edge 30 and outboard of the distal end of the initial increased wall thickness zone 50 to the blade tip 44, has a span or stand length, L_T, which is 240 mm.

[0026] Relative differences in airfoil outer wall or side wall thickness zone 50 along the trailing edge 30 are defined as follows, referring to the blade cross section of FIGs. 4-6. In the initial increased outer wall thickness zone 50, the airfoil outer wall thickness ¾, excluding adjoining hub fillet 46 thickness, may be approximately forty to sixty percent (40-60%) greater along the trailing edge 30 (in the stand length LH) than the comparable greatest wall thickness, T_T, anywhere else along the trailing edge 30 from outboard that zone 50 all the way to the blade tip 44 (i.e., in the span L_T). In some embodiments, the outer wall thickness T_H is constant in the L_H stand length portion of zone 50. To assist greater comprehension of the respective local relative airfoil outer wall or side wall 36 thicknesses in these described embodiments, FIGs. 4-6 show the actual wall thickness T_H in the thickness zone 50 in the cross section taken through cooling passage 62, which is delimited by the outer wall exterior surface 38 and the interior surface 40. The thickness TH and the corresponding thickness Ττ for the distal tip stand length LT region (outboard of the initial increased thickness zone 50) is shown in the dotted lines inboard of the actual outer wall exterior surfaces 38. As noted above, corresponding airfoil outer wall or side wall thickness in the transition zone (in the span or stand length L_x) may transition from that of the adjoining greatest increased wall proximate the hub platform 24, i.e., in the proximal portion within the thickness zone 50, to the corresponding distal end of the initial increased thickness zone's thickness of the outboard adjoining outer wall thickness Ττ in the stand length LT. Actual absolute thickness or any thickness variations in the trailing edge 30 portion outer wall thickness Ττ, outboard the thickness zone 50 out to the blade tip 44, along the span or stand length L_T is constant, or varied, such as by tapering to thinner thickness toward the end of the blade tip 44.

[0027] Initial increased trailing edge outer wall thickness in the zone 50 directly changes the relative casting solidification rates of the airfoil trailing edge 30 and the blade platform 24 mass in a direction that results in improved grain structure from the casting process. This improvement increases the material's rupture capability as confirmed by micro structural evaluations and elevated temperature creep rupture testing. These comparisons of a thickened TE 30 blade casting, with the zone 50, as compared to nominal thickness blade castings, confirmed the effects of this feature upon the material's rupture capability. Grain size affects the crack initiation and propagation in the concentrated stress areas. Initial increased airfoil outer wall 38 thickness in the zone 50 directly influences these formations in a positive direction, by changing relative solidification rate to reduce the formation of fine grains that result in reduced ductility and corresponding concentrated creep rupture capability. [0028] Figure 4 shows a desirable shape for a high efficiency thin trailing edge 30. As previously noted, the airfoil 26 has an initial increased airfoil side wall or outer wall 36 thickness zone 50 that is proximate the trailing edge 30. This initial increased airfoil side wall or outer wall 36 thickness zone 50 may be produced during the casting and solidification process of manufacturing. As shown in FIG. 5, the initial increased outer wall thickness zone 50 has a proximal end that initiates at the airfoil proximal of the hub end 42 and runs upwardly along the blade stand for approximately eight to ten percent (8-10%) of the total airfoil span, or stand length dimension, which is denoted by L_H.

[0029] Figure 4 shows a trailing edge with a minimal thickness within the thickness zone 50. Figure 8 illustrates steps within a method for manufacturing a turbine blade 20 for a turbine engine. During the casting process, an additional material section 70 is added to the initial increased airfoil side wall or outer wall 36 thickness zone 50. Figure 5 shows a method to slow down the solidification rate of the casting by introducing a larger wall thickness designated by T_¾ versus the wall thickness Ττ that is desired and shown in Figure 4. Figures 6 and 7 are different embodiments of how additional material sections 70 may be added to the trailing edge to provide initial additional thickness in the airfoil side wall or outer wall 36 thickness zone 50 for the casting process. This additional material section 70 may be of any length and size, with Figures 6 and 7 shown as two of several examples of the shape and size of the additional material section 70. A mold for the turbine blade may be produced with a portion of the mold removed for this additional material section 70. An initial increased airfoil side wall or outer wall 36 thickness zone 50 may be produced with this mold. A liquid material, generally liquid metal, but not limited to, is then poured into the mold and cured, and then solidified. The liquid material made from a superalloy, metal, metals, or the like. The mold is a negative for the eventual blade 20. Breaking of the mold once the solidification is complete reveals the blade 20 as casted. Once the blade 20 is casted and solidified, the additional material section 70 can clearly be seen as is shown in Figures 6 and 7 providing an as-cast surface. The additional material section 70, such as the "wing" example in Figure 6 or the extra layer in Figure 7, may be machined off as is shown in Figures 6 and 7 reducing the overall thickness in the airfoil side wall or outer wall 36 thickness zone 50 providing a final decreased airfoil side wall or outer wall 36 thickness zone 50 and final machined surface to a more desired thickness Ττ for aerodynamic purposes such as is shown in Figure 4. The shapes in Figures 6 and 7 are shown as examples, however, are not exhaustive as to shape of the additional material section 70. [0030] The casting may have features at the trailing edge that are locally increased in thermal mass and slow down the solidification rate at this location. An example is for the trailing edge made twice, three times, or the like of a typical thickness. After the blade has been cast and solidified, the extra metal may be removed by machining if necessary until desired thickness is reached while leaving a thin trailing edge with desirable crystal structure.

[0031] Initial increased trailing edge outer wall thickness in the zone 50 directly changes the relative casting solidification rates of the airfoil trailing edge 30 and the blade platform 24 mass in a direction that results in improved grain structure from the casting process. The addition as casted, and the removal of the additional material section 70 provides a method to control the grain structure by introducing a larger wall thickness at the trailing edge and then machining it off back to a high- aerodynamic efficiency shape shown in Figure 4, with a thickness of T_T. This improvement increases the metal's rupture capability as confirmed by micro structural evaluations and elevated temperature concentrated creep rupture testing. Grain size affects the crack initiation and propagation in the concentrated stress areas. The addition of a manufacturing step to machine off a percentage of the trailing edge outer wall thickness in the zone 50 improves the aerodynamic characteristics of the trailing edge, and therefore the blade.

[0032] The Figures and description used above are applicable to cooled blades, however, the same principles and concepts of these embodiments are also applicable to an un-cooled solid blade that has no cooling passages. [0033] While specific embodiments have been described in detail, those with ordinary skill in the art will appreciate that various modifications and alternative to those details could be developed in light of the overall teachings of the disclosure. Accordingly, the particular arrangements disclosed are meant to be illustrative only and not limiting as to the scope of the invention, which is to be given the full breadth of the appended claims, and any and all equivalents thereof.

[0034] Although various embodiments that incorporate the invention have been shown and described in detail herein, others can readily devise many other varied embodiments that still incorporate the claimed invention. The invention is not limited m 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 invention is capable of other embodiments and of being practiced or of being carried out in various ways. In addition, 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 ' ncluding," "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 ai'e used broadly and encompass direct and indirect mountings, connections, supports, and couplings.

Claims

CLAIMS What is claimed is:

1. A turbine blade (20) for a turbine engine comprising:

a hub, including a blade platform (24);

an elongated airfoil portion (26), having:

an outer wall (36) delimiting a pressure side, a suction side, a leading edge (28), and a trailing edge (30) on an exterior surface thereof, an airfoil interior on an interior surface thereof, and an outer wall thickness between the respective interior and exterior surfaces;

a proximal end (42) of the outer wall coupled to the blade platform from the leading edge to the trailing edge;

a distal end (44) of the outer wall defining a blade tip; and an airfoil span (L) defined between the proximal and distal ends thereof;

an initial increased airfoil outer wall thickness zone (50), having a zone proximal end adjoining the blade platform and a zone distal end, wherein the zone outer wall thickness greater along the trailing edge for a portion of total airfoil span most proximate the hub than comparable greatest outer wall thickness anywhere else along the trailing edge from outboard the zone distal end all the way to the blade tip, with outer wall thickness at the zone distal end transitioning to that of the adjoining outboard outer wall thickness, wherein a temporary additional material section (70) is added to the outer wall thickness zone; and

a final decreased airfoil outer wall thickness zone (50), wherein the zone outer wall thickness is approximately the same as the total airfoil span most proximate the hub as the comparable greatest outer wall thickness anywhere else along the trailing edge from outboard the zone distal end all the way to the blade tip, wherein the temporary additional material section (70) is removed.

2. The turbine blade (20) of claim 1, further comprising a hub fillet circumscribing and joined to the airfoil outer wall exterior surface at the proximal end thereof, and joined to the blade platform.

3. The turbine blade of claims 1 or 2, further comprising a plurality of pedestals spanning the airfoil interior, having first and second ends coupled to respective corresponding opposed interior surfaces of the outer wall pressure and suction sides proximate the trailing edge, the pedestals oriented along at least part of the airfoil stand length between the proximal and distal ends thereof

4. The turbine blade of any of claims 1 through 3, the airfoil interior proximate the trailing edge defining a cooling gap between the opposed outer wall interior surfaces, for passage of cooling fluid there through, the pedestals spanning the cooling gap.

5. A method for manufacturing a turbine blade for a turbine engine, comprising:

producing a mold for casting, the mold to comprise a turbine blade comprising;

a hub, including a blade platform (24);

an elongated airfoil portion (26), having:

an outer wall delimiting a pressure side, a suction side, a leading edge, and a trailing edge on an exterior surface thereof, an airfoil interior on an interior surface thereof, and an outer wall thickness between the respective interior and exterior surfaces;

a proximal end of the outer wall coupled to the blade platform from the leading edge to the trailing edge;

a distal end (44) of the outer wall (36) defining a blade tip; and airfoil span (L) defined between the proximal and distal ends thereof; removing a portion of the mold to create an additional material section (70) along an airfoil outer wall thickness zone (50), having a zone proximal end adjoining the blade platform and a zone distal end, wherein the zone outer wall thickness is greater along the trailing edge for a portion of total airfoil span most proximate the hub than comparable greatest outer wall thickness anywhere else along the trailing edge from outboard the zone distal end all the way to the blade tip, with outer wall thickness at the zone distal end transitioning to that of the adjoining outboard outer wall thickness, wherein additional material is attached to the outer wall thickness zone; and pouring a liquid material into the mold;

solidifying liquid material;

breaking mold revealing the blade component; and

machining off the additional material section (70) if necessary from the airfoil outer wall thickness zone (50), until desired thickness is reached.