Office Action 1: Ex Parte Quayle Action (2025-06-02)

CLAIMS:
What is claimed is:

1\. A gas turbine engine, comprising:an engine core configured to generate a redline exhaust gas temperature (EGT) in a range of 988 Celsius (0C) to 1120 0C and having a redline core speed (CS) in a range of 306 Hertz to 353 Hertz, the engine core extending along an engine centerline and including:a compressor section;
a combustor; and
a turbine section, the turbine section including a blade assembly rotatable about the engine centerline, the blade assembly comprising:
a platform having an upper surface and a lower surface;
an airfoil extending from the upper surface of the platform, the airfoil having an exterior surface, the exterior surface defining a pressure side and a suction side, the airfoil including a leading-edge and a trailing-edge;
a shank extending from the lower surface, the shank having a base defining a base plane;
a middle inlet passage located within the shank, the middle inlet passage having:
a first cooling conduit portion at a first plane radially spaced 0.0 167 meters from the base plane, the first cooling conduit portion having a first cross-sectional area, the first cooling portion defining a first bounding circle, the first cooling portion having:
a first point at a first 12:30 o'clock position on the first bounding circle; and
a second point at a first 1:30 o'clock position on the first bounding circle, the first point and the second point defining a first line;

a second cooling conduit portion at a second plane radially spaced 0.0197 meters from the base plane, the second cooling conduit portion having a second cross-sectional area;
a third cooling conduit portion at a third plane radially spaced 0.0215 meters from the base plane, the third cooling conduit portion having a third cross-sectional area;
a fourth cooling conduit portion at a fourth plane radially spaced 0.0234 meters from the base plane, the fourth cooling conduit portion having a fourth cross-sectional area;
a fifth cooling conduit portion at a fifth plane radially spaced 0.0245 meters from the base plane, the fifth cooling conduit portion having a fifth cross-sectional area, the fifth cooling portion defining a second bounding circle, the second cooling portion having:a third point at a second 12:30 o'clock position on the second bounding circle; and a fourth point at a third 1:30 o'clock position on the second bounding circle, the third point and the fourth point defining a second line;wherein the middle inlet passage defines a maximum normalized cooling conduit portion area (Amaxeen 1.30 and 2.10, Amaxed by:A = m axSVG {EMBED HERE} 0.3430.41ein the first line and 2 the second line define a degree of change of the cooling conduit portions along the suction side (DCCPASS) ranging between 21.0 and 33.0 degrees (0); and a plurality of cooling conduits located within the airfoil, the plurality of cooling conduits including a quantity (N1) of cooling conduits fed by the middle inlet passage, wherein the quantity is 2, and
wherein:25.3530.75\*(DCCPASS (o) 2SVG {EMBED HERE} 0.2332.863;149.922.

2\. The gas turbine engine of claim 1,further including a plurality of inlet passages located within the shank, the plurality of inlet passages including the middle inlet passage.

3\. The gas turbine engine of claim 2, wherein each of the inlet passages extends between the base and one or more of the plurality of cooling conduits.

4\. The gas turbine engine of claim 2, wherein the plurality of inlet passages further includes a leading-edge inlet passage and a trailing-edge inlet passage.

5\. The gas turbine engine of claim 1, wherein the shank is configured as a dovetail.

6\. The gas turbine engine of claim 1, wherein the blade assembly is a stage one blade assembly of a high pressure turbine of the turbine section.

7\. The gas turbine engine of claim 1, wherein the quantity (N1) of cooling conduits fed by the middle inlet passage includes:a forward serpentine cooling conduit in the airfoil; and
an aft serpentine cooling conduit in the airfoil.

8\. The gas turbine engine of claim 7, wherein the middle inlet passage is fluidly coupled to the forward serpentine cooling conduit and the aft serpentine cooling conduit at a splitter, the splitter including:a first cooling conduit portion at a sixth plane radially spaced 0.449 meters from the base plane, the first cooling conduit portion associated with the forward serpentine cooling conduit, the first cooling conduit portion having a first major axis;
and a second cooling conduit portion at the sixth plane, the second cooling conduit portion associated with the aft serpentine cooling conduit, the second cooling conduit portion having a second major axis, wherein:

the first major axis and the second major axis define an L-shape.
9\. The gas turbine engine of claim 7, wherein the quantity is a first quantity, the forward serpentine cooling conduit defines a second quantity of cooling conduit portions spanning pressure surface to suction surface, the cooling conduit portions taken at a sixth plane radially spaced 0.449 meters from the base plane, and the second quantity is three.

10\. A blade assembly for a gas turbine engine having an engine core configured to generate a redline exhaust gas temperature (EGT) in a range of 988 Celsius (4C) to 1120 0C and having a redline core speed (CS) in a range of 306 Hertz to 353 Hertz, the blade assembly to be connected to the engine core and rotatable about an engine centerline of the engine core, the blade assembly comprising:a platform having an upper surface and a lower surface;
an airfoil extending from the upper surface of the platform, the airfoil defining an exterior surface, the exterior surface defining a pressure side and a suction side, the airfoil including a leading-edge and a trailing-edge;
a shank extending from the lower surface, the shank having a base defining a base plane;
a middle inlet passage located within the shank, the middle inlet passage having:
a first cooling conduit portion at a first plane radially spaced 0.0167 meters from the base plane, the first cooling conduit portion having a first cross-sectional area, the first cooling conduit portion having a first cross- sectional area, the first cooling portion defining a first bounding circle, the first cooling portion having:
a first point at a first 12:30 o'clock position on the first bounding circle; and
a second point at a first 1:30 o'clock position on the first bounding circle, the first point and the second point defining a first line;

a second cooling conduit portion at a second plane radially spaced 0.0197 meters from the base plane, the second cooling conduit portion having a second cross-sectional area;
a third cooling conduit portion at a third plane radially spaced 0.02 15 meters from the base plane, the third cooling conduit portion having a third cross-sectional area;
a fourth cooling conduit portion at a fourth plane radially spaced 0.0234 meters from the base plane, the fourth cooling conduit portion having a fourth cross-sectional area;
a fifth cooling conduit portion at a fifth plane radially spaced 0.0245 meters from the base plane, the fifth cooling conduit portion having a fifth cross-sectional area, the fifth cooling portion defining a second bounding circle, the second cooling portion having:a third point at a second 12:30 o'clock position on the second bounding circle; and
a fourth point at a third 1:30 o'clock position on the second bounding circle, the third point and the fourth point defining a second line;
wherein the middle inlet passage defines a maximum normalized cooling conduit portion area (Amaxing between 1.30 and 2.10, Amax defined by:{sAmax = max {Aein the first line and Ai=2 the second line define a degree of change of the cooling conduit portions along the suction side (DCCPASS) ranging between 21.0 and 33.0 degrees (°); and a plurality of cooling conduits located within the airfoil, the plurality of cooling conduits including a quantity (N1) of cooling conduits fed by the middle inlet passage, wherein the quantity is 2, and wherein:25.353 0.75 \* (C zcc\*(Anss\*SVG {EMBED HERE} 0.2372.513149922

11\. The blade assembly of claim 10, further including a plurality of inlet passages located within the shank, the plurality of inlet passages including the middle inlet passage.

12\. The gas turbine engine of claim 11, wherein each of the inlet passages extends between the base and one or more of the plurality of cooling conduits.

13\. The gas turbine engine of claim ll, wherein the plurality of inlet passages further includes a leading-edge inlet passage and a trailing-edge inlet passage.

14\. The blade assembly of claim 10, wherein the shank is configured as a dovetaiL

15\. The blade assembly of claim 10, wherein the blade assembly is a stage one blade assembly of a high pressure turbine of the gas turbine engine.

16\. The blade assembly of claim 10, wherein the quantity (N1) of cooling conduits fed by the middle inlet passage includes:a forward serpentine cooling conduit in the airfoil; and
an aft serpentine cooling conduit in the airfoil.

17\. The blade assembly of claim 16, wherein the middle inlet passage is fluidly coupled to the forward serpentine cooling conduit and the aft serpentine cooling conduit at a splitter, the splitter including:a first cooling conduit portion at a sixth plane radially spaced 0.449 meters from the base plane, the first cooling conduit portion associated with the forward serpentine cooling conduit, the first cooling conduit portion having a first major axis;
and a second cooling conduit portion at the sixth plane, the second cooling conduit portion associated with the aft serpentine cooling conduit, the second cooling conduit portion having a second major axis, wherein:the first major axis and the second major axis define an L-shape.

18\. The blade assembly of claim 16, wherein the quantity is a first quantity, the forward serpentine cooling conduit defines a second quantity of cooling conduit portions spanning pressure surface to suction surface, the cooling conduit portions taken at a sixth plane radially spaced 0.449 meters from the base plane, and the second quantity is three.

OFFICE ACTION:
DETAILED ACTIONNotice of Pre-AIA  or AIA  Status
The present application, filed on or after March 16, 2013, is being examined under the first inventor to file provisions of the AIA .
Claim Objections
Claim 1 is objected to because of the following informalities:
the limitation “the second cooling portion” at lines 41-42 should be “the fifth cooling portion”.
the limitation “Amaxeen” at line 49 should be “Amax between”
the limitation “Amaxed” at line 50 should be “Amax defined”
the limitation “ein” at line 51 should be “wherein”
the limitation “the cooling conduit portions” at line 52 should be “the first, second, third, fourth and fifth cooling conduit portions”.
Claim 9, is objected to because the limitation “spanning pressure surface to suction surface” should be “spanning from the pressure side to the suction side”.
Claim 10 is objected to because:
the limitation “the first cooling conduit portion having a first cross sectional area” at lines 16-17 is a duplicate typographical error and should be deleted.
the limitation “the second cooling portion” at line 36 should be “the fifth cooling portion”
the limitation “Amax ing” at line 43 should be “Amax”
the limitation “ein” at line 45 should be “wherein”
the limitation “cooling conduit portions” at line 46 should be ““the first, second, third, fourth and fifth cooling conduit portions”.
Claim 18 is objected to because the limitation “spanning pressure surface to suction surface” should be “spanning from the pressure side to the suction side”.Appropriate correction is required.
Allowable Subject Matter
Claims 1-18 are otherwise allowed.
The following is an examiner’s statement of reasons for allowance: the prior art does not teach or fairly suggest the claimed cooling conduit portions in the middle inlet passage such that a maximum normalized cooling conduit potion area as claimed presents a DCCPASS as claimed satisfying the function recited in the last lines of claims 1 and 10. Prior art attached herein are considered pertinent but not anticipatory or otherwise obvious.
Any comments considered necessary by applicant must be submitted no later than the payment of the issue fee and, to avoid processing delays, should preferably accompany the issue fee.  Such submissions should be clearly labeled “Comments on Statement of Reasons for Allowance.”
Conclusion
This application is in condition for allowance except for the following formal matters:
The objections identified above.
Prosecution on the merits is closed in accordance with the practice under Ex parte Quayle, 25 USPQ 74, 453 O.G. 213, (Comm’r Pat. 1935).
A shortened statutory period for reply to this action is set to expire TWO (2) MONTHS from the mailing date of this letter. Extensions of time may be granted under  37 CFR 1.136 but in no case can any extension carry the date for reply to this Office action beyond the maximum period of SIX MONTHS set by statute (35 U.S.C. 133).


NEXT OFFICE ACTION: DETAILED ACTION
Notice of Pre-AIA  or AIA  Status
The present application, filed on or after March 16, 2013, is being examined under the first inventor to file provisions of the AIA .
Allowable Subject Matter
Claims 1-18 are allowed.
The following is an examiner’s statement of reasons for allowance: The Examiner notes the counterpart opinion from the office. The specification at Table 2, pp. 17-18, appears to show the DCCPASS angles and the maximum normalized cooling conduit portion area. Additionally, the counterpart opinion appears to indicate the claimed Amax and DCCPASS angle at the claimed portions is a mere design choice in view of the prior art. However, the prior art does not appear to discuss a basis for the design choice nor the area and the DCCPASS angles being a result effective variable to warrant a rejection. Therefore, the claims remain allowable as indicated previously. The Applicant having corrected all remaining issues, the case should proceed to allowance.
Any comments considered necessary by applicant must be submitted no later than the payment of the issue fee and, to avoid processing delays, should preferably accompany the issue fee.  Such submissions should be clearly labeled “Comments on Statement of Reasons for Allowance.”
Conclusion

CITED REFERENCE SUMMARIES: Ref US4501053A: Claims:
A method for making a coolable rotor blade for an axial flow rotary machine which has a spanwisely extending spar having a root section having cooling air passages extending therethrough and an airfoil section and which has a shell for defining the airfoil surface of the rotor blade, which includes the steps of: forming a spanwisely extending spar with a projection on the spar which extends chordwisely about the spar; joining the shell to the airfoil section of the spar and to the projection on the spar such that the shell extends telescopically over and against a portion of the projection on the spar; and forming a platform about the spar and the shell such that the platform extends over the projection on the spar and over and against the portion of the shell extending over the projection on the spar to trap the shell between the platform and the projection on the spar.
forming a spanwisely extending spar with a projection on the spar which extends chordwisely about the spar;
joining the shell to the airfoil section of the spar and to the projection on the spar such that the shell extends telescopically over and against a portion of the projection on the spar; and
forming a platform about the spar and the shell such that the platform extends over the projection on the spar and over and against the portion of the shell extending over the projection on the spar to trap the shell between the platform and the projection on the spar.
The method of making a coolable rotor blade of claim 1 wherein the step of joining the shell to the airfoil section of the spar includes spacing the shell from the root of the spar and the step of forming a platform about the spar and shell includes forming the platform such that it extends between the root section and the shell for blocking the leakage of cooling air from the interior of the blade between the shell and the spar to the exterior of the blade.
The method of making a coolable rotor blade of claim 2 wherein the step of forming the platform about the spar and shell includes the step of mechanically attaching the platform to the spar.
The method of making a coolable rotor blade of claim 1 claim 2, claim 3, or claim 4 wherein the step of forming a spanwisely extending spar includes the step of forming the spar of one material and the step of forming the platform includes the step of forming the platform from a different material.
The method of making a coolable rotor blade of claim 3 wherein the step of forming a spanwisely extending spar includes the step of forming the spar from a first material, wherein the method includes the step of forming a shell from a second material and wherein the step of forming the platform from a different material includes the step of forming the platform from a third material.
The method of making a coolable rotor blade of claim 8 wherein the step of forming a shell includes the step of forming a suction sidewall and a pressure sidewall, wherein the step of joining the shell to the airfoil section of the spar includes the step of attaching the pressure sidewall to the spar and the suction sidewall to the spar and wherein the method further includes the step of attaching the pressure sidewall and the suction sidewall one to the other at the leading edge region and wherein the method further includes the step of attaching a tip cap to the tip region of the shell.
The method of making a coolable rotor blade of claim 7 wherein the step of forming a platform about the spar and the shell includes the step of forming the platform of a multilayer fiber reinforced ceramic material.
The method of forming a coolable rotor blade of claim 7 wherein the step of forming a platform about the spar and the shell includes the step of forming the platform of a metallic alloy.
The method of making a coolable rotor blade of claim 8 wherein the step of forming a platform about the spar and the shell takes place before the step of attaching the tip cap and includes the step of forming the platform of a silicon carbide fiber reinforced ceramic by molding ceramic constituents in powder form to the shape of a platform about the shell and the spar wherein the constituents are in the glassy state (noncrystalline form) and are combined in the powder state with the silicon carbide fibers by hot pressed consolidation followed by heating of the constituents at a time and a temperature sufficient to transform the noncrystalline glassy state components into the crystalline ceramic state by controlled nucleation and growth and wherein during the step of forming the platform the method includes the step of flowing a cooling fluid through the cooling passages in the spar and through the shell.
The method of making a coolable rotor blade of claim 8 wherein the step of forming a platform about the spar and the shell includes the step of forming a first portion of the platform of a fiber reinforced glass composite with the matrix (glass) in the glassy state and forming a second portion of the platform of a fiber reinforced glass composite with the matrix (glass) in the glassy state and disposing the two portions about the spar and shell and subsequently heating the two pieces together and applying pressure to the two pieces to convert the matrix (glass) to the crystalline state (ceramic).
The method of making a coolable rotor blade of claim 11 wherein the step of forming a platform about the spar and shell includes the step of inserting a dowel of the third material such that it extends between the first portion and the second portion and the step of heating the fiber reinforced glass composite portions with the matrix (glass) in the glassy state includes the step of heating the dowel and the portions together and applying pressure until the heat and pressure converts the matrix of each piece to the crystalline state and bonds the portions together and bonds the dowel to both portions.
The method of making a coolable rotor blade as claimed in claim 10 wherein the subsequent heating of the fiber reinforced glass composite with the matrix (glass) in the glassy state takes place at a temperature in excess of two thousand degrees Fahrenheit (2000Â° F.).
The method of making a coolable rotor blade as claimed in claim 11 wherein the subsequent heating of the fiber reinforced glass composite with the matrix (glass) in the glassy state takes place at a temperature in excess of two thousand degrees Fahrenheit (2000Â° F.).
The method of making a coolable rotor blade as claimed in claim 12 wherein the subsequent heating of the fiber reinforced glass composite with the matrix (glass) in the glassy state takes place at a temperature in excess of two thousand degrees Fahrenheit (2000Â° F.).
Ref US5993155A: Claims:
A cooled gas-turbine blade of solid construction comprising: a blade having a lateral surface, a blade thickness, a blade root, a blade tip, a cooling-medium feed, a cooling medium discharge, and cooling passages which run in the interior of the blade through which cooling passages cooling medium can flow from said blade root through said blade to cool said blade, wherein said cooling passages have a relatively small cross section in relation to said blade thickness and are arranged so as to run close to said blade lateral surface, said cooling-medium feed and cooling-medium discharge being formed in said blade root, wherein each of said cooling passages comprises a U-shaped branch, and wherein each of said cooling passages is fluidly separated from other cooling passages from said cooling-medium feed to said cooling-medium discharge.
a blade having a lateral surface, a blade thickness, a blade root, a blade tip, a cooling-medium feed, a cooling medium discharge, and cooling passages which run in the interior of the blade through which cooling passages cooling medium can flow from said blade root through said blade to cool said blade, wherein said cooling passages have a relatively small cross section in relation to said blade thickness and are arranged so as to run close to said blade lateral surface, said cooling-medium feed and cooling-medium discharge being formed in said blade root, wherein each of said cooling passages comprises a U-shaped branch, and wherein each of said cooling passages is fluidly separated from other cooling passages from said cooling-medium feed to said cooling-medium discharge.
The gas-turbine blade as claimed in claim 1, wherein said cooling passage branches each include two stems, and said cooling passage branches are nested one inside the other so that each stem of a branch is essentially the neighbor of another branch.
A cooled gas-turbine blade of solid construction comprising: a blade having a lateral surface, a blade thickness, a blade root, a blade tip, a cooling-medium feed, a cooling medium discharge, and cooling passages which run in the interior of the blade through which cooling passages cooling medium can flow from said blade root through said blade to cool said blade, wherein said cooling passages have a relatively small cross section in relation to said blade thickness and are arranged so as to run close to said blade lateral surface, said cooling-medium feed and cooling-medium discharge being formed in said blade root; and a bypass line in said blade root fluidly communicating said cooling-medium feed and said cooling medium discharge so that when cooling medium is caused to flow from said cooling-medium feed, through said cooling passages, and out said cooling-medium discharge, particles entrained in said cooling medium are not allowed to pass into said cooling passages.
a blade having a lateral surface, a blade thickness, a blade root, a blade tip, a cooling-medium feed, a cooling medium discharge, and cooling passages which run in the interior of the blade through which cooling passages cooling medium can flow from said blade root through said blade to cool said blade, wherein said cooling passages have a relatively small cross section in relation to said blade thickness and are arranged so as to run close to said blade lateral surface, said cooling-medium feed and cooling-medium discharge being formed in said blade root; and
a bypass line in said blade root fluidly communicating said cooling-medium feed and said cooling medium discharge so that when cooling medium is caused to flow from said cooling-medium feed, through said cooling passages, and out said cooling-medium discharge, particles entrained in said cooling medium are not allowed to pass into said cooling passages.
Ref US20030219338A1: Claims:
A method for manufacturing a blade for a gas turbine engine, wherein the blade includes an airfoil, a platform, a shank, and a dovetail, the platform extending between the airfoil and the shank, the shank extending between the dovetail and the platform, the dovetail including at least one tang for securing the blade within the engine, said method comprising: defining a cooling cavity in the blade that extends through the airfoil the platform, the shank, and the dovetail, wherein the portion of the cavity defined within the dovetail includes a root passage portion having a first width, and a transition portion that extends between the root passage and the portion of the cavity defined within the shank, and wherein the portion of the cavity defined within the shank has a second width that is larger than the root passage first width; and coating at least a portion of an inner surface of the blade that defines the cooling cavity with a layer of an oxidation resistant environmental coating.
defining a cooling cavity in the blade that extends through the airfoil the platform, the shank, and the dovetail, wherein the portion of the cavity defined within the dovetail includes a root passage portion having a first width, and a transition portion that extends between the root passage and the portion of the cavity defined within the shank, and wherein the portion of the cavity defined within the shank has a second width that is larger than the root passage first width; and
coating at least a portion of an inner surface of the blade that defines the cooling cavity with a layer of an oxidation resistant environmental coating.
A method in accordance with claim 1 wherein said defining a cooling cavity further comprises defining the cooling cavity within the dovetail such that the root passage first width is substantially constant within the root passage, and such that the transition portion is tapered between the root passage and the portion of the cavity defined within the shank, such that a width of the transition portion is variable within the transition portion.
A method in accordance with claim 2 wherein said defining a cooling cavity further comprises defining the cooling cavity such that an interface between the dovetail transition portion and the shank portion forms an arcuate shape that defines a portion of the cooling cavity.
A method in accordance with claim 2 wherein said coating at least a portion of an inner surface of the blade further comprises coating at least a portion of the inner surface of the cooling cavity within the dovetail with a coating having a thickness greater than 0.001 inches.
A method in accordance with claim 1 wherein said coating at least a portion of an inner surface of the blade further comprises coating an inner surface of the cooling cavity within the dovetail to facilitate reducing life cycle fatigue cracking within the dovetail.
A blade for a gas turbine engine, said blade comprising: a platform; a shank extending from said platform; a dovetail extending between an end of said blade and said shank for mounting said blade within the gas turbine engine, said dovetail comprising at least one tang; an airfoil comprising a first sidewall and a second sidewall extending in radial span between said platform and a blade tip; and a cooling cavity defined within said blade by said dovetail, said shank, said platform, and said airfoil, said cooling cavity comprising a dovetail portion defined within said dovetail, a shank portion defined within said shank and said platform, and an airfoil portion defined within said airfoil, said shank portion coupled in flow communication between said airfoil portion and said dovetail portion, said dovetail portion comprising a root passage and a transition passage, said root passage comprising a first width, said shank portion comprising a second width larger than said root passage first width, said transition passage coupled between said root passage and said shank portion.
a shank extending from said platform;
a dovetail extending between an end of said blade and said shank for mounting said blade within the gas turbine engine, said dovetail comprising at least one tang;
an airfoil comprising a first sidewall and a second sidewall extending in radial span between said platform and a blade tip; and
a cooling cavity defined within said blade by said dovetail, said shank, said platform, and said airfoil, said cooling cavity comprising a dovetail portion defined within said dovetail, a shank portion defined within said shank and said platform, and an airfoil portion defined within said airfoil, said shank portion coupled in flow communication between said airfoil portion and said dovetail portion, said dovetail portion comprising a root passage and a transition passage, said root passage comprising a first width, said shank portion comprising a second width larger than said root passage first width, said transition passage coupled between said root passage and said shank portion.
A blade in accordance with claim 6 wherein said cooling cavity root passage first width measured between a pressure side of said cooling cavity and a suction side of said cooling cavity, said root passage first width substantially constant within said root passage.
A blade in accordance with claim 6 wherein said cooling cavity shank passage second width measured between a pressure side of said cooling cavity and a suction side of said cooling cavity, an interface of said transition passage and said shank portion is arcuate.
A blade in accordance with claim 8 wherein said cooling cavity interface facilitates reducing operating stresses induced within said blade dovetail.
A blade in accordance with claim 6 wherein said dovetail further comprises an inner surface defining said cooling cavity dovetail portion, said dovetail inner surface coated with an oxidation resistant environmental coating.
A blade in accordance with claim 6 wherein said dovetail further comprises an inner surface defining said cooling cavity dovetail portion, said dovetail inner surface coated with an oxidation resistant environmental coating having a thickness greater than 0.001 inches.
A blade in accordance with claim 6 wherein said cooling cavity configured to facilitate reducing dovetail low cycle fatigue cracking.
A gas turbine engine comprising a plurality of blades, each said blade comprising an airfoil, a shank, and a platform extending therebetween, each said blade further comprising a cooling cavity, and a dovetail comprising at least one tang and configured to secure each said blade within said engine, said shank extending between said platform and said dovetail, said cooling cavity defined within said airfoil, said platform, said shank, and said dovetail, said cooling cavity comprising a dovetail portion, a shank portion, and an airfoil portion coupled in flow communication, said cooling cavity dovetail portion comprising a root passage comprising a first width, and a transition passage, said cooling cavity shank portion comprising a second width that is larger than said root passage first width, said cooling cavity transition passage tapering between said root passage and said shank portion.
A gas turbine engine in accordance with claim 13 wherein said cooling cavity root passage first width measured between a pressure side and a suction side of said cooling cavity, said cooling cavity shank portion second width measured between said cooling cavity pressure and suction sides, said root passage first width substantially constant within said root passage.
A gas turbine engine in accordance with claim 14 wherein an interface between said cooling cavity transition passage and said cooling cavity shank portion forms a radius.
A gas turbine engine in accordance with claim 14 wherein said cooling cavity coated with an oxidation resistant environmental coating having a thickness configured to reduce low cycle fatigue of each said blade.
A gas turbine engine in accordance with claim 14 wherein at least a portion of said cooling cavity coated with an oxidation resistant environmental coating having a thickness greater than 0.001 inches.
A gas turbine engine in accordance with claim 14 wherein at least a portion of said cooling cavity dovetail portion coated with an oxidation resistant environmental coating having a thickness greater than 0.001 inches.
A gas turbine engine in accordance with claim 14 wherein each said cooling cavity facilitates reducing operating stresses induced within each said blade dovetail.
A gas turbine engine in accordance with claim 14 wherein said cooling cavity is configured to facilitate reducing dovetail low cycle fatigue cracking.
Ref US20160237833A1: Claims:
A turbine blade comprising: an airfoil and a fir tree root, the fir tree root having a lengthwise direction, a crosswise direction extending between two lateral sides of the fir tree root, and a span direction extending from a root base towards an airfoil tip, the fir tree root having at least one longitudinal groove arranged on each lateral side and extending along and defining the lengthwise direction, the fir tree root having a root thickness measured between the two lateral sides, said width varying along the span direction, the fir tree root having at least two channels extending in the span direction, each of said channels having a chan
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