Office Action 1: Notice of Allowance (2025-06-17)

CLAIMS:
CLAIMS What is claimed is:
1\. A turbine engine comprising: an engine core defining an engine centerline; sets of airfoils comprising a set of part-span guide vanes and a set of fan blades, the sets airfoils circumferentially arranged about the engine centerline, an airfoil of the sets of airfoils comprising: a main body portion extending chordwise between a body leading edge and a trailing edge; and a leading edge protector receiving at least a portion of the body leading edge of the main body portion, the leading edge protector extending chordwise from a protector leading edge towards the trailing edge of the main body portion to define a leading length (LL); wherein the main body portion and the leading edge protector together define an exterior surface of the airfoil, wherein the protector leading edge is different from and forward of the body leading edge, and wherein a chord length (CL) is defined as extending chordwise between the protector leading edge and the trailing edge of the main body portion; and wherein the leading length (LL) is related to the chord length (CL) by an airfoil protection factor (APF) = LLCL and a first APF (APF1) for set of fan blades is greater than or equal to 0.07 and less than or equal to 0.67 (0.07 ≤ APF1 ≤ 0.67) and a second APF (APF2) for the set of part-span guide vanes is greater than or equal to 0.04 and less than or equal to 10 (0.04 ≤ APF2 ≤ 0.10).
2\. The turbine engine of claim 1, wherein the set of fan blades includes a composite airfoil.
3\. The turbine engine of claim 2, wherein the composite airfoil includes a polymer matrix composite (PMC).
4\. The turbine engine of claim 3, wherein the leading edge protector is a metallic leading edge protector.
5\. The turbine engine of claim 1, wherein the leading edge protector is a metallic leading edge protector.
6\. The turbine engine of claim 5, wherein the set of part-span guide vanes include metallic airfoils.
7\. The turbine engine of claim 1, wherein the set of part-span guide vanes are part-span inlet guide vanes.
8\. The turbine engine of claim 7, wherein the part-span inlet guide vanes are cantilevered.
9\. The turbine engine of claim 7, wherein the set of fan blades has a first number of airfoils and the set of part-span inlet guide vanes has a second number of airfoils and the first number is less than the second number.
10\. The turbine engine of claim 9, wherein the second number is two to three times greater than the first number.
11\. The turbine engine of claim 1, wherein the airfoil extends spanwise between a root and a tip to define a span length and wherein the APF is measured from 20% to 80% of the span length.
12\. A turbine engine comprising: an engine core defining an engine centerline; a set of fan blades circumferentially arranged about the engine centerline, an airfoil of the sets of fan blades comprising: a main body portion extending chordwise between a body leading edge and a trailing edge of the main body portion; and a leading edge protector receiving at least a portion of the body leading edge of the main body portion, the leading edge protector extending chordwise from a protector leading edge towards the trailing edge of the main body portion to define a leading length (LL); wherein the main body portion and the leading edge protector together define an exterior surface of the airfoil, wherein the protector leading edge is different from and forward of the body leading edge, and wherein a chord length (CL) is defined as extending chordwise between the protector leading edge and the trailing edge of the main body portion; and wherein the airfoil extends spanwise between a root and a tip to define a span length and wherein the leading length (LL) and the chord length (CL) are measured from 20% to 80% of the span length, wherein the leading length (LL) is related to the chord length (CL) by an airfoil protection factor (APF) = LLCL and the APF for the set of fan blades is greater than or equal to 0.07 and less than or equal to 0.67 (0.07 ≤ APF1 ≤ 0.67).
13\. The turbine engine of claim 12, wherein the LL ranges from 5 cm to 20 cm (5 cm ≤ LL ≤ 20 cm).
14\. The turbine engine of claim 13, wherein the CL ranges from 30 cm to 71 cm (30 cm ≤ CL ≤ 71 cm).
15\. The turbine engine of claim 12, wherein the LL ranges from 7.5 cm to 16.5 cm (7.5 cm ≤ LL ≤ 16.5 cm).
16\. The turbine engine of claim 12, wherein the CL ranges from 38 cm to 65 cm (38 cm ≤ CL ≤ 65 cm).
17\. The turbine engine of claim 12, wherein the main body portion is a composite main body portion.
18\. The turbine engine of claim 17, wherein the leading edge protector is a metallic leading edge protector.
19\. The turbine engine of claim 12, wherein the turbine engine provides thrust in a range of greater than or equal to 35,000 pounds of thrust to less than or equal to 70,000 pounds of thrust.
20\. The turbine engine of claim 12, further comprising a nacelle surrounding and at least partially enclosing the set of fan blades.

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 .
EXAMINER'S AMENDMENT
An examiner’s amendment to the record appears below. Should the changes and/or additions be unacceptable to applicant, an amendment may be filed as provided by 37 CFR 1.312. To ensure consideration of such an amendment, it MUST be submitted no later than the payment of the issue fee.
Authorization for this examiner’s amendment was given in an interview with Natalie Armstrong on May 29, 2025.
The application has been amended as follows:
Claim 1. (Currently Amended) A turbine engine comprising:
an engine core defining an engine centerline;
sets of airfoils comprising a set of part-span guide vanes and a set of fan blades, the sets of airfoils circumferentially arranged about the engine centerline, an airfoil of the sets of airfoils comprising:
a main body portion extending chordwise between a body leading edge and a trailing edge; and
a leading edge protector receiving at least a portion of the body leading edge of the main body portion, the leading edge protector extending chordwise from a protector leading edge towards the trailing edge of the main body portion to define a leading length (LL);
wherein the main body portion and the leading edge protector together define an exterior surface of the airfoil, wherein the protector leading edge is different from and forward of the body leading edge, and wherein a chord length (CL) is defined as extending chordwise between the protector leading edge and the trailing edge of the main body portion; and
wherein the leading length (LL) is related to the chord length (CL) by an airfoil protection factor (APF) =
(
L
L
)
(
C
L
)
and a first APF (APF1) for the set of fan blades is greater than or equal to 0.07 and less than or equal to 0.67 (0.07 ≤ APF1 ≤ 0.67) and a second APF (APF2) for the set of part-span guide vanes is greater than or equal to 0.04 and less than or equal to 10 (0.04 ≤ APF2 ≤ 0.10).
Claim 12. (Currently Amended) A turbine engine comprising:
an engine core defining an engine centerline;
a set of fan blades circumferentially arranged about the engine centerline, an airfoil of the set of fan blades comprising:
a main body portion extending chordwise between a body leading edge and a trailing edge of the main body portion; and
a leading edge protector receiving at least a portion of the body leading edge of the main body portion, the leading edge protector extending chordwise from a protector leading edge towards the trailing edge of the main body portion to define a leading length (LL);
wherein the main body portion and the leading edge protector together define an exterior surface of the airfoil, wherein the protector leading edge is different from and forward of the body leading edge, and wherein a chord length (CL) is defined as extending chordwise between the protector leading edge and the trailing edge of the main body portion; and
wherein the airfoil extends spanwise between a root and a tip to define a span length and wherein the leading length (LL) and the chord length (CL) are measured from 20% to 80% of the span length, wherein the leading length (LL) is related to the chord length (CL) by an airfoil protection factor (APF) =
(
L
L
)
(
C
L
)
and the APF for the set of fan blades is greater than or equal to 0.07 and less than or equal to 0.67 (0.07 ≤ APF1 ≤ 0.67).
Reasons for Allowance
The following is an examiner’s statement of reasons for allowance: Examiner notes the closest prior art to that of claim 1 is US 10,724,435 to Kroger et al. which discloses similar primary and upstream subsidiary airfoils, though not including any protective sheath/shield and no APF’s.  US 2021/0388726 to Churcher discloses the closest art in the form of a shield/sheath applied to leading edges, and US 8,814,527 to Huth et al. discloses a ratio of surface area of an airfoil shield being desirable.  Examiner notes; however, the combined features do not clarify for the claimed range of APF as claimed, given the equation of surface area and length ratios as disclosed by Huth, above, would not account for the additional thickness of the protective sheaths as projecting forward of the leading edge as disclosed and amended, and thus the values provided by Huth, cannot be confirmed to apply within the range of the APF provided in claim 1 (and in claim 12).  Given the subsidiary airfoils upstream can be very small, in particular, (with regards to claim 1 APF 2) (See Table 4 of the parent application 18489533) the examiner notes, Huth when applied at say 50% coverage per Huth ( a value of 1 in Huth being half coverage half not) the thickness of the leading edge protector as compared to the chord of the subsidiary airfoil may be large enough to remove the value arrived at outside of the APF ratio as claimed between LL and CL of the fan airfoils, which cut off closely to 50%, at 0.7, and more critically for the static guide vanes.
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 US10724435B2: Claims:
A gas turbine engine defining an axial direction and a radial direction, the gas turbine engine comprising: a turbomachine; a fan rotatable by the turbomachine, the fan comprising a plurality of fan blades; an outer nacelle surrounding the plurality of fan blades; and a plurality of part-span inlet guide vanes cantilevered from the outer nacelle at a location forward of the plurality of fan blades along the axial direction, each of the plurality of part-span inlet guide vanes defining an inner end along the radial direction and being unconnected with an adjacent part-span inlet guide vane at the inner end, and each of the plurality of part-span inlet guide vanes defining a first swirl angle at a trailing edge proximate the inner end and a second swirl angle at the trailing edge proximate an outer end, and wherein the first swirl angle and the second swirl angle differ to provide a reduction in flow separation and/or shock losses.
a fan rotatable by the turbomachine, the fan comprising a plurality of fan blades;
an outer nacelle surrounding the plurality of fan blades; and
a plurality of part-span inlet guide vanes cantilevered from the outer nacelle at a location forward of the plurality of fan blades along the axial direction, each of the plurality of part-span inlet guide vanes defining an inner end along the radial direction and being unconnected with an adjacent part-span inlet guide vane at the inner end, and each of the plurality of part-span inlet guide vanes defining a first swirl angle at a trailing edge proximate the inner end and a second swirl angle at the trailing edge proximate an outer end, and wherein the first swirl angle and the second swirl angle differ to provide a reduction in flow separation and/or shock losses.
The gas turbine engine of claim 1 , wherein the plurality of part-span inlet guide vanes each define the outer end along the radial direction, and wherein each of the plurality of part-span inlet guide vanes is connected to the outer nacelle at the outer end.
The gas turbine engine of claim 1 , wherein each of the plurality of part-span inlet guide vanes is completely supported by the outer nacelle at the respective outer end.
The gas turbine engine of claim 1 , wherein the plurality of part-span inlet guide vanes define a solidity between about 0.5 and about 1.5.
The gas turbine engine of claim 1 , wherein the plurality of part-span inlet guide vanes define a solidity between about 0.9 and 1.0.
The gas turbine engine of claim 1 , wherein each of the plurality of part-span inlet guide vanes is configured as a variable part-span inlet guide vane.
The gas turbine engine of claim 1 , wherein the plurality of part-span inlet guide vanes includes between twenty part-span inlet guide vanes and fifty part-span inlet guide vanes.
The gas turbine engine of claim 1 , wherein the plurality of part-span inlet guide vanes includes between thirty part-span inlet guide vanes and forty-five part-span inlet guide vanes.
The gas turbine engine of claim 1 , wherein each of the plurality of part-span inlet guide vanes defines a leading edge, the trailing edge, and a maximum swirl angle, and wherein the maximum swirl angle of each of the plurality of part-span inlet guide vanes at the trailing edge is between five degrees and thirty-five degrees.
The gas turbine engine of claim 1 , wherein each of the plurality of part-span inlet guide vanes defines a leading edge, the trailing edge, and a maximum swirl angle, and wherein the maximum swirl angle of each of the plurality of part-span inlet guide vanes at the trailing edge is between twelve degrees and twenty-five degrees.
The gas turbine engine of claim 1 , wherein the second swirl angle is greater than the first swirl angle.
The gas turbine engine of claim 1 , wherein the inner end of each of the plurality of inlet guide vanes is movable generally along the radial direction between an extended position and a retracted position.
The gas turbine engine of claim 1 , wherein the gas turbine engine further defines a circumferential direction, and wherein the plurality of part-span inlet guide vanes are spaced substantially evenly along the circumferential direction.
The gas turbine engine of claim 1 , wherein the gas turbine engine further defines a circumferential direction, and wherein the plurality of part-span inlet guide vanes are spaced unevenly along the circumferential direction.
The gas turbine engine of claim 1 , wherein the plurality of part-span inlet guide vanes each define a longitudinal axis, wherein the gas turbine engine further defines a circumferential direction, wherein the radial and circumferential directions together define a reference plane, and wherein the longitudinal axis of each of the plurality of part-span inlet guide vanes intersects the reference plane and defines a sweep angle with the reference plane between about ten degrees and about thirty degrees.
The gas turbine engine of claim 1 , wherein the trailing edge of each of the plurality of part-span inlet guide vanes defines a non-linear sculpted shape.
The gas turbine engine of claim 1 , further comprising: a compensation air supply assembly in airflow communication with a high pressure air source to receive a compensation airflow, wherein the plurality of part-span inlet guide vanes each defines the trailing edge and a trailing edge opening, wherein the trailing edge opening is in airflow communication with the compensation air supply assembly to receive the composition airflow and provide such compensation airflow through the trailing edge opening during operation of the gas turbine engine.
a compensation air supply assembly in airflow communication with a high pressure air source to receive a compensation airflow, wherein the plurality of part-span inlet guide vanes each defines the trailing edge and a trailing edge opening, wherein the trailing edge opening is in airflow communication with the compensation air supply assembly to receive the composition airflow and provide such compensation airflow through the trailing edge opening during operation of the gas turbine engine.
A gas turbine engine defining an axial direction and a radial direction, the gas turbine engine comprising: a turbomachine; a fan rotatable by the turbomachine, the fan comprising a plurality of fan blades; an outer nacelle surrounding the plurality of fan blades; and a plurality of part-span inlet guide vanes cantilevered from the outer nacelle at a location forward of the plurality of fan blades along the axial direction, each of the plurality of inlet guide vanes defining an outer end along the radial direction and being completely supported by the outer nacelle at the respective outer end, and each of the plurality of part-span inlet guide vanes defining a first swirl angle at a trailing edge proximate an inner end and a second swirl angle at the trailing edge proximate the outer end, and wherein the first swirl angle and the second swirl angle differ to provide a reduction in flow separation and/or shock losses.
a plurality of part-span inlet guide vanes cantilevered from the outer nacelle at a location forward of the plurality of fan blades along the axial direction, each of the plurality of inlet guide vanes defining an outer end along the radial direction and being completely supported by the outer nacelle at the respective outer end, and each of the plurality of part-span inlet guide vanes defining a first swirl angle at a trailing edge proximate an inner end and a second swirl angle at the trailing edge proximate the outer end, and wherein the first swirl angle and the second swirl angle differ to provide a reduction in flow separation and/or shock losses.
The gas turbine engine of claim 18 , wherein each of the plurality of part-span inlet guide vanes is unconnected with an adjacent part-span inlet guide vane at the inner end.
The gas turbine engine of claim 18 , wherein the plurality of part-span inlet guide vanes define a solidity between about 0.5 and about 1.5.
Ref US20210388726A1: Claims:
A method of manufacturing a component for a gas turbine engine, the method comprising: applying a thermoplastic polymer sheet over a composite body for the component; applying a shield over part of the composite body, the shield terminating at an end which overlies the thermoplastic polymer sheet and defines an interface between shielded and unshielded regions of the component; and pressing the shield into the thermoplastic polymer sheet so that the thermoplastic polymer sheet deforms around the end of the shield, such that an exterior profile of the component at the interface between the shielded and unshielded regions is flush.
applying a thermoplastic polymer sheet over a composite body for the component;
applying a shield over part of the composite body, the shield terminating at an end which overlies the thermoplastic polymer sheet and defines an interface between shielded and unshielded regions of the component; and
pressing the shield into the thermoplastic polymer sheet so that the thermoplastic polymer sheet deforms around the end of the shield, such that an exterior profile of the component at the interface between the shielded and unshielded regions is flush.
The method according to claim 1 comprising: pressing the shield into the thermoplastic polymer sheet so that the thermoplastic polymer sheet deforms around the end of the shield such that the thermoplastic polymer sheet has a discontinuous thickness at the end of the shield.
pressing the shield into the thermoplastic polymer sheet so that the thermoplastic polymer sheet deforms around the end of the shield such that the thermoplastic polymer sheet has a discontinuous thickness at the end of the shield.
The method according to claim 2 comprising: pressing the shield into the thermoplastic polymer sheet so that the thermoplastic polymer sheet deforms around the end of the shield such that the thickness of the thermoplastic polymer sheet ( 43 ) between the shield and the composite body is reduced relative to the thickness of the thermoplastic polymer sheet in the unshielded region.
pressing the shield into the thermoplastic polymer sheet so that the thermoplastic polymer sheet deforms around the end of the shield such that the thickness of the thermoplastic polymer sheet ( 43 ) between the shield and the composite body is reduced relative to the thickness of the thermoplastic polymer sheet in the unshielded region.
The method according to claim 1 comprising: pressing the shield into the thermoplastic polymer sheet using a tool which defines a profile for the component.
pressing the shield into the thermoplastic polymer sheet using a tool which defines a profile for the component.
The method according to claim 1 , wherein the thermoplastic polymer sheet comprises thermoplastic polyurethane.
The method according to claim 1 , wherein the shield is a metal shield.
The method according to claim 1 , wherein the end of the shield is chamfered.
The method according to claim 1 further comprising: applying adhesive between the thermoplastic polymer sheet and the composite body; and/or applying adhesive between the shield and the composite body and/or the thermoplastic polymer sheet.
applying adhesive between the thermoplastic polymer sheet and the composite body; and/or
applying adhesive between the shield and the composite body and/or the thermoplastic polymer sheet.
The method according to claim 8 , wherein the adhesive is supported by a scrim material.
The method according to claim 1 , wherein the component is a fan blade or a guide vane and the composite body is a composite fan blade body or a composite guide vane body, wherein the composite fan blade body or the composite guide vane body comprises a pressure surface and a suction surface, each extending between a leading edge and a trailing edge, and wherein the method comprises: applying the shield over the leading edge and/or the trailing edge.
applying the shield over the leading edge and/or the trailing edge.
The method according to claim 1 , wherein the component is a nose cone, the composite body is a composite nose cone body, and the method comprises: applying the shield over an apex of the composite nose cone body.
applying the shield over an apex of the composite nose cone body.
A component for a gas turbine engine manufactured according to a method comprising: applying a thermoplastic polymer sheet over a composite body for the component; applying a shield over part of the composite body, the shield terminating at an end which overlies the thermoplastic polymer sheet and defines an interface between shielded and unshielded regions of the component; and pressing the shield into the thermoplastic polymer sheet so that the thermoplastic polymer sheet deforms around the end of the shield, such that an exterior profile of the component at the interface between the shielded and unshielded regions is flush.
A component for a gas turbine engine, the component comprising: a composite body; a thermoplastic polymer layer extending over the composite body; and at least one shield extending over part of the composite body, the or each shield terminating at an end which overlies the thermoplastic polymer layer and defines an interface between a respective shielded region and an unshielded region of the component, an exterior profile of the component at the or each interface between the shielded and unshielded regions being flush; wherein at the end of the or each shield, the thermoplastic polymer layer has a discontinuous thickness, the thickness of the thermoplastic polymer layer between the shield and the composite body in the respective shielded region being lower than the thickness of the thermoplastic polymer layer in the unshielded region; and wherein the thermoplastic polymer layer covers a remainder of the composite body not covered by any of the at least one shield.
a composite body;
a thermoplastic polymer layer extending over the composite body; and
at least one shield extending over part of the composite body, the or each shield terminating at an end which overlies the thermoplastic polymer layer and defines an interface between a respective shielded region and an unshielded region of the component, an exterior profile of the component at the or each interface between the shielded and unshielded regions being flush;
wherein at the end of the or each shield, the thermoplastic polymer layer has a discontinuous thickness, the thickness of the thermoplastic polymer layer between the shield and the composite body in the respective shielded region being lower than the thickness of the thermoplastic polymer layer in the unshielded region; and
wherein the thermoplastic polymer layer covers a remainder of the composite body not covered by any of the at least one shield.
The component according to claim 13 , wherein the component is a fan blade or a guide vane and the composite body is a composite fan blade body or a composite guide vane body, the fan blade or the guide vane comprising a pressure surface and a suction surface, each extending between a leading edge and a trailing edge, and the or each shield extending over the leading edge and/or the trailing edge.
The component according to claim 13 , wherein the component is a nose cone and the composite body is a composite nose cone body, a shield of the at least one shield extending over an apex of the nose cone body.
Ref US8814527B2: Claims:
An airfoil assembly comprising: an airfoil body extending between a leading edge, a trailing edge, a suction side, and a pressure side; a first sheath including a first plate on the suction side of the airfoil body and a second plate on the pressure side of the airfoil body that meet at a seam at a nose near the leading edge, the first sheath having an outer side and an inner side that forms a cavity for receiving at least a portion of the airfoil body, the inner side being bonded to the leading edge of the airfoil body, and the first sheath being formed from a titanium-based metallic material; and a second sheath bonded to the outer side of the first sheath.
an airfoil body extending between a leading edge, a trailing edge, a suction side, and a pressure side;
a first sheath including a first plate on the suction side of the airfoil body and a second plate on the pressure side of the airfoil body that meet at a seam at a nose near the leading edge, the first sheath having an outer side and an inner side that forms a cavity for receiving at least a portion of the airfoil body, the inner side being bonded to the leading edge of the airfoil body, and the first sheath being formed from a titanium-based metallic material; and
a second sheath bonded to the outer side of the first sheath.
The airfoil assembly as recited in claim 1 , wherein a wall thickness of the first sheath tapers from a nose at the leading edge of the airfoil body to a rearward end of the first sheath.
The airfoil assembly as recited in claim 1 , wherein the second sheath is formed of a nickel-based metallic material.
The airfoil assembly as recited in claim 1 , further comprising an adhesive between the first sheath and the airfoil body.
The airfoil assembly as recited in claim 1 , wherein the airfoil body is a composite material comprising a polymer matrix and reinforcement fibers distributed within the polymer matrix.
The airfoil assembly as recited in claim 1 , wherein the airfoil body includes surface portions that are covered by the first sheath and other surface portions that are not covered by the first sheath, with a ratio of the covered portion to the uncovered portion that is about 0.1-1.
The airfoil assembly as recited in claim 1 , wherein the inner side of the first sheath is in direct contact with the airfoil body, and the second sheath is adhesively bonded directly to the outer side of the first sheath.
The airfoil assembly as recited in claim 1 , wherein the second sheath is seam-free.
The airfoil assembly as recited in claim 1 , wherein the second sheath has a free outer surface.
The airfoil assembly as recited in claim 1 , wherein the second sheath has a free outer surface at a nose.
The airfoil assembly as recited in claim 1 , wherein the second sheath is formed of a nickel-based metallic material, the first sheath providing ductility that absorbs energy from an impact and the second sheath providing a lower ductility and a higher hardness for erosion.
An airfoil assembly comprising: an airfoil body extending between a leading edge, a trailing edge, a suction side, and a pressure side; a first sheath having an outer side and an inner side that forms a cavity for receiving at least a portion of the airfoil body, the inner side being bonded to the leading edge of the airfoil body, the first sheath being formed from a titanium-based metallic material, the first sheath including a first plate on the suction side of the airfoil body and a second plate on the pressure side of the airfoil body that meet at a nose near the leading edge, wherein the first plate and the second plate include respective notches that abut against the leading edge of the airfoil body, the first plate and the second plate each having a uniform wall thickness from the respective notches to a tip of the nose, and wherein the nose of the first sheath extends forward relative to the leading edge of the airfoil body and includes a seam where the first plate and the second plate meet; and a second sheath bonded to the outer side of the first sheath.
a first sheath having an outer side and an inner side that forms a cavity for receiving at least a portion of the airfoil body, the inner side being bonded to the leading edge of the airfoil body, the first sheath being formed from a titanium-based metallic material, the first sheath including a first plate on the suction side of the airfoil body and a second plate on the pressure side o
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