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 .
Continued Examination Under 37 CFR 1.114
A request for continued examination under 37 CFR 1.114, including the fee set forth in 37 CFR 1.17(e), was filed in this application after allowance or after an Office action under Ex Parte Quayle, 25 USPQ 74, 453 O.G. 213 (Comm'r Pat. 1935). Since this application is eligible for continued examination under 37 CFR 1.114, and the fee set forth in 37 CFR 1.17(e) has been timely paid, prosecution in this application has been reopened pursuant to 37 CFR 1.114. Applicant's submission filed on 08/26/2026 has been entered.
Claim Status
Claims 1-4, 9-14, 17, 19-24 are pending:
Claims 23-24 are new
Claims 5-8, 15-16, 18 are canceled
Response to Arguments
No arguments were provided.
Double Patenting
The nonstatutory double patenting rejection is based on a judicially created doctrine grounded in public policy (a policy reflected in the statute) so as to prevent the unjustified or improper timewise extension of the “right to exclude” granted by a patent and to prevent possible harassment by multiple assignees. A nonstatutory double patenting rejection is appropriate where the conflicting claims are not identical, but at least one examined application claim is not patentably distinct from the reference claim(s) because the examined application claim is either anticipated by, or would have been obvious over, the reference claim(s). See, e.g., In re Berg, 140 F.3d 1428, 46 USPQ2d 1226 (Fed. Cir. 1998); In re Goodman, 11 F.3d 1046, 29 USPQ2d 2010 (Fed. Cir. 1993); In re Longi, 759 F.2d 887, 225 USPQ 645 (Fed. Cir. 1985); In re Van Ornum, 686 F.2d 937, 214 USPQ 761 (CCPA 1982); In re Vogel, 422 F.2d 438, 164 USPQ 619 (CCPA 1970); In re Thorington, 418 F.2d 528, 163 USPQ 644 (CCPA 1969).
A timely filed terminal disclaimer in compliance with 37 CFR 1.321(c) or 1.321(d) may be used to overcome an actual or provisional rejection based on nonstatutory double patenting provided the reference application or patent either is shown to be commonly owned with the examined application, or claims an invention made as a result of activities undertaken within the scope of a joint research agreement. See MPEP § 717.02 for applications subject to examination under the first inventor to file provisions of the AIA as explained in MPEP § 2159. See MPEP § 2146 et seq. for applications not subject to examination under the first inventor to file provisions of the AIA . A terminal disclaimer must be signed in compliance with 37 CFR 1.321(b).
The filing of a terminal disclaimer by itself is not a complete reply to a nonstatutory double patenting (NSDP) rejection. A complete reply requires that the terminal disclaimer be accompanied by a reply requesting reconsideration of the prior Office action. Even where the NSDP rejection is provisional the reply must be complete. See MPEP § 804, subsection I.B.1. For a reply to a non-final Office action, see 37 CFR 1.111(a). For a reply to final Office action, see 37 CFR 1.113(c). A request for reconsideration while not provided for in 37 CFR 1.113(c) may be filed after final for consideration. See MPEP §§ 706.07(e) and 714.13.
The USPTO Internet website contains terminal disclaimer forms which may be used. Please visit www.uspto.gov/patent/patents-forms. The actual filing date of the application in which the form is filed determines what form (e.g., PTO/SB/25, PTO/SB/26, PTO/AIA /25, or PTO/AIA /26) should be used. A web-based eTerminal Disclaimer may be filled out completely online using web-screens. An eTerminal Disclaimer that meets all requirements is auto-processed and approved immediately upon submission. For more information about eTerminal Disclaimers, refer to www.uspto.gov/patents/apply/applying-online/eterminal-disclaimer.
Claim 12, 13, 17 provisionally rejected on the ground of nonstatutory double patenting as being unpatentable over claim 13, 15 of copending Application No. 19/424,772 (reference application), as indicated below. Although the claims at issue are not identical, they are not patentably distinct from each other.
This is a provisional nonstatutory double patenting rejection because the patentably indistinct claims have not in fact been patented.
Please see the comparison table below; the underlined recitations indicate a feature not claimed in the other patent document.
Instant claim 12 is rejected relative to both claims 13/15 of copending application 19/424,772.
Copending claim 9 discloses all of Instant Claim 12 except for the claimed gear ratios, however each of copending claims 13/15 disclose gear ratios within the range claimed in instant claim 9.
Instant claim 13 is rejected relative to copending claim 13/15
Instant claim 17 is rejected relative to copending claim 13/15
Instant Application 19/362,542
Copending Application 19/424,772
Claim 12
Claim 9
A gas turbine engine defining a radial direction, the gas turbine engine comprising:
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
an unducted fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFanLE and a trailing edge fan radius RFanTE, and the fan defining a leading edge hub radius RHUb_LE and a trailing edge hub radius RHUb_TE, the gas turbine engine defining a bypass ratio equal to a mass flowrate of an airflow from the unducted fan over the turbomachine to a mass flowrate of an airflow from the unducted fan through the inlet to the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100; and
a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the unducted fan, wherein the reduction gearbox defines a gear ratio greater than 2 and less than 14;
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Opening Ratio (FLTOR) greater than or equal to 1.03 and less than or equal to 1.5, the FLTOR being equal to:
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An aircraft comprising:
a fuselage;
an airfoil extending from the fuselage, the airfoil having an airfoil section defining an effective quarter chord point (QC);
two or more unducted gas turbine engine defining a radial direction, each unducted gas turbine engine comprising:
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
a fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFan_LE and a trailing edge fan radius RFan_TE, and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, the gas turbine engine defining a bypass ratio equal to a mass flowrate of an airflow from the fan over the turbomachine to a mass flowrate of an airflow from the fan through the inlet to the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100; and
a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan;
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Opening Ratio (FLTOR) greater than or equal to 1.03 and less than or equal to 1.5, the FLTOR being equal to:
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wherein each of the two or more unducted gas turbine engines is mounted relative to the airfoil section on a high pressure side thereof, the unducted gas turbine engine having a centerline (CL), a plurality of blades arranged in a forward array and a plurality of blades arranged in a rearward array, wherein only one of the forward and rearward array of blades are rotating blades that include the fan, the rotating blades defining a maximum outer diameter (D);a point (P) located at an intersection of the CL and a line HP perpendicular to the CL that passes through an axial midpoint between the rearward trailing edge at a root of a blade of the rearward array and a forward trailing edge at a root of the blade of the forward array when the forward leading edge and rearward leading edge of the respective blades are aligned with each other; and an ellipse origin positioning line (EOR) having a length (EORL) extending from the QC to an ellipse origin (OR) at an angle 0 measured positive in a counter-clockwise direction when the high pressure side of the airfoil section is below the airfoil section, and measured positive in a clockwise direction when the high pressure side of the airfoil section is above the airfoil section, when viewed looking for an outboard position towards an inboard position; wherein the P of the unducted gas turbine engine is located within a first ellipse having a first major axis length (1MajAL) and a first minor axis length (1MinAL) with a first ellipse origin defined by EORL/D of 0.938 and 0 of 253.6°, and where 1MajAL/D is 2.8 and 1MinAL/D is 1.7.
Claim 12 provisionally rejected on the ground of nonstatutory double patenting as being unpatentable over claim 14, 16, 17 of copending Application No. 19/532,138 (reference application), as indicated below. Although the claims at issue are not identical, they are not patentably distinct from each other.
This is a provisional nonstatutory double patenting rejection because the patentably indistinct claims have not in fact been patented.
Please see the comparison table below; the underlined recitations indicate a feature not claimed in the other patent document.
Instant claim 12 is rejected relative to claims 14/16/17 of copending application 19/532,138.
Copending claim 10 discloses all of Instant Claim 12 except for the claimed gear ratios, however each of copending claims 14/16/17 disclose gear ratios within the range claimed in instant claim 12.
Instant Application 19/362,542
Copending Application 19/532,138
Claim 12
Claim 10
A gas turbine engine defining a radial direction, the gas turbine engine comprising:
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
an unducted fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFanLE and a trailing edge fan radius RFanTE, and the fan defining a leading edge hub radius RHUb_LE and a trailing edge hub radius RHUb_TE, the gas turbine engine defining a bypass ratio equal to a mass flowrate of an airflow from the unducted fan over the turbomachine to a mass flowrate of an airflow from the unducted fan through the inlet to the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100; and
a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the unducted fan, wherein the reduction gearbox defines a gear ratio greater than 2 and less than 14;
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Opening Ratio (FLTOR) greater than or equal to 1.03 and less than or equal to 1.5, the FLTOR being equal to:
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A gas turbine engine defining a radial direction, the gas turbine engine comprising:
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
a fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFan_LE and a trailing edge fan radius RFan_TE, and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, the gas turbine engine defining a bypass ratio equal to a mass flowrate of an airflow from the fan over the turbomachine to a mass flowrate of an airflow from the fan through the inlet to the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100; and
a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan;
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Opening Ratio (FLTOR) greater than or equal to 1.03 and less than or equal to 1.5, the FLTOR being equal to:
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wherein the composite material comprises a fiber preform architecture comprising: a plurality of fiber bands, each of the plurality of fiber bands placed one at a time, at a predetermined position and orientation, to generate an interlocking pattern between fiber bands, wherein each of the plurality of fiber bands comprise a plurality of fiber tows, positioned side by side at a predetermined spacing to define one or more gaps between each tow in the fiber band, wherein the plurality of fiber tows and bands are interwoven in an in-plane and out-of-plane orientation by interleaving each of the plurality of fiber tows and bands with one or more of the plurality of fiber tows and bands previously laid down and not in a common plane to fill the one or more gaps and define a uniformly covered multi-layered assembly, and wherein the plurality of fiber bands include three or more different orientation angles.
Claim 1, 12 provisionally rejected on the ground of nonstatutory double patenting as being unpatentable over claim 6, 10, 11 of copending Application No. 19/531,849 (reference application), as indicated below. Although the claims at issue are not identical, they are not patentably distinct from each other.
This is a provisional nonstatutory double patenting rejection because the patentably indistinct claims have not in fact been patented.
Please see the comparison table below; the underlined recitations indicate a feature not claimed in the other patent document.
Instant claim 1 is rejected relative to claim 6
Copending claim 6 discloses all of Instant Claim 1 except for a nacelle surrounding the fan, however gas turbines without a fan nacelle (open rotor) and gas turbines with a fan nacelle (closed rotor) are well known alternatives in the art.
Instant claim 12 is rejected relative to claims 10/11 of copending application 19/531,849
Copending claim 7 discloses all of Instant Claim 12 except for the claimed gear ratios, however each of copending claims 10/11 disclose gear ratios within the range claimed in instant claim 12.
Instant Application 19/362,542
Copending Application 19/531,849
Claim 1
Claim 6 (1+6)
A gas turbine engine defining a radial direction, the gas turbine engine comprising:
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
a fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFanLE and a trailing edge fan radius and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, wherein the leading edge fan radius RFan LE is greater than or equal to 20 inches and less than or equal to 85 inches, the gas turbine engine defining a bypass ratio equal to a mass flowrate of an airflow from the fan over the turbomachine to a mass flowrate of an airflow from the fan through the inlet to the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100;
an outer nacelle at least partially surrounding the fan; and
a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan, the reduction gearbox defining a gear ratio greater than or equal to 2 and less than or equal to 4;
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8, the FLTCF being equal to:
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A gas turbine engine defining a radial direction, the gas turbine engine comprising:
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
a fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFan_LE and a trailing edge fan radius RFan_TE, and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, the gas turbine engine defining a bypass ratio equal to a mass flowrate of an airflow from the fan over the turbomachine to a mass flowrate of an airflow from the fan through the inlet to the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100; and
a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan;
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8, the FLTCF being equal to:
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and wherein the fan blade comprises a composite portion extending chordwise between a composite leading edge and a trailing edge; and a leading edge protector receiving at least a portion of the composite leading edge of the composite portion, the leading edge protector extending chordwise from a leading edge around the composite portion on both a pressure side of the composite portion and a suction side of the composite portion, and wherein the fan blade has a straight line chord length (CL), a leading edge protector chord length on the pressure side (LLP), a leading edge protector chord length on the suction side (LLS), the fan having a number of blades (Nb), wherein the gas turbine engine has a bypass ratio greater than 13:1, and wherein 0.25 ≤ (LLP/ LLS) * CL / Nb ≤ 2.40.
Claim 6
The gas turbine engine of claim 1, wherein the leading edge fan radius RFan_LE is greater than or equal to 35 inches and less than or equal to 50 inches, wherein the fan defines a fan blade count greater than or equal to 12 and less than or equal to 23, and wherein the reduction gearbox defines a gear ratio between 2:1 and 4:1.
Claim 12
Claim 7
A gas turbine engine defining a radial direction, the gas turbine engine comprising:
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
an unducted fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFanLE and a trailing edge fan radius RFanTE, and the fan defining a leading edge hub radius RHUb_LE and a trailing edge hub radius RHUb_TE, the gas turbine engine defining a bypass ratio equal to a mass flowrate of an airflow from the unducted fan over the turbomachine to a mass flowrate of an airflow from the unducted fan through the inlet to the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100; and
a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the unducted fan, wherein the reduction gearbox defines a gear ratio greater than 2 and less than 14;
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Opening Ratio (FLTOR) greater than or equal to 1.03 and less than or equal to 1.5, the FLTOR being equal to:
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A gas turbine engine defining a radial direction, the gas turbine engine comprising:
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
a fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFan_LE and a trailing edge fan radius RFan_TE, and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, the gas turbine engine defining a bypass ratio equal to a mass flowrate of an airflow from the fan over the turbomachine to a mass flowrate of an airflow from the fan through the inlet to the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100; and
a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan;
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8, the FLTCF being equal to:
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and wherein the fan blade comprises a composite portion extending chordwise between a composite leading edge and a trailing edge; and a leading edge protector receiving at least a portion of the composite leading edge of the composite portion, the leading edge protector extending chordwise from a leading edge around the composite portion on both a pressure side of the composite portion and a suction side of the composite portion, and wherein the fan blade has a straight line chord length (CL), a leading edge protector chord length on the pressure side (LLP), a leading edge protector chord length on the suction side (LLS), the fan having a number of blades (Nb), wherein the gas turbine engine has a bypass ratio greater than 13:1, and wherein 0.25 ≤ (LLP/ LLS) * CL / Nb ≤ 2.40.
Claim 7
The gas turbine engine of claim 1,
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Opening Ratio (FLTOR) greater than or equal to 1.03 and less than or equal to 1.5, the FLTOR being equal to:
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Claim 1 provisionally rejected on the ground of nonstatutory double patenting as being unpatentable over claim 8 of copending Application No. 19/545,392 (reference application), as indicated below. Although the claims at issue are not identical, they are not patentably distinct from each other.
This is a provisional nonstatutory double patenting rejection because the patentably indistinct claims have not in fact been patented.
Please see the comparison table below; the underlined recitations indicate a feature not claimed in the other patent document.
Instant claim 1 is rejected relative to claim 8
Copending claim 8 discloses all of Instant Claim 1 except for a nacelle surrounding the fan, however gas turbines without a fan nacelle (open rotor) and gas turbines with a fan nacelle (closed rotor) are well known alternatives in the art.
Instant Application 19/362,542
Copending Application 19/545,392
Claim 1
Claim 8 (1+6+8)
A gas turbine engine defining a radial direction, the gas turbine engine comprising:
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
a fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFanLE and a trailing edge fan radius and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, wherein the leading edge fan radius RFan LE is greater than or equal to 20 inches and less than or equal to 85 inches, the gas turbine engine defining a bypass ratio equal to a mass flowrate of an airflow from the fan over the turbomachine to a mass flowrate of an airflow from the fan through the inlet to the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100;
an outer nacelle at least partially surrounding the fan; and
a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan, the reduction gearbox defining a gear ratio greater than or equal to 2 and less than or equal to 4;
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8, the FLTCF being equal to:
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A gas turbine engine defining a radial direction, the gas turbine engine comprising:
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
a fan having a plurality of fan blades formed of a composite material, each respective fan blade of the plurality of fan blades defining a leading edge fan radius RFan LE and a trailing edge fan radius RFnand the fan defining a leading edge hub radius Rub_LE and a trailing edge hub radiusthe gas turbine engine defining a bypass ratio equal to a mass flowrate of an airflow from the fan over the turbomachine to a mass flowrate of an airflow from the fan through the inlet to the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100; and a
reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan;
wherein: the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8, the FLTCF being equal to:
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the fan includes a first VPF parameter and a second VPF parameter;the first VPF parameter is defined by a fan blade radius ratio (RR) divided by a fan pressure ratio (FPR) at a static sea-level takeoff operating condition;the second VPF parameter is defined by a bearing spanwise force (F Span) at a redline operating condition measured in pounds force divided by a fan area (F Area) measured in square inches; andthe first VPF parameter is within a range of 0.1 to 0.25 and the second VPF parameter is within a range of 2-30 lbf/in2, or the first VPF parameter is within a range of 0.1 to 0.4 and the second VPF parameter is within a range of 5.25-30 lbf/in2.
Claim 6
The gas turbine engine of claim 1, further comprising: an outer nacelle surrounding at least in part the fan.
Claim 8
The gas turbine engine of claim 6, wherein the leading edge fan radius RFanLE is greater than or equal to 65 inches and less than or equal to 85 inches, and wherein the fan defines a fan blade count greater than or equal to 5 and less than or equal to 15.
Claim 1 provisionally rejected on the ground of nonstatutory double patenting as being unpatentable over claim 9 of copending Application No. 19/545,325 (reference application), as indicated below. Although the claims at issue are not identical, they are not patentably distinct from each other.
This is a provisional nonstatutory double patenting rejection because the patentably indistinct claims have not in fact been patented.
Please see the comparison table below; the underlined recitations indicate a feature not claimed in the other patent document.
Instant claim 1 is rejected relative to claim 9
Copending claim 9 discloses all of Instant Claim 1 except for a nacelle surrounding the fan, however gas turbines without a fan nacelle (open rotor) and gas turbines with a fan nacelle (closed rotor) are well known alternatives in the art.
Instant Application 19/362,542
Copending Application 19/545,325
Claim 1
Claim 9 (1+9)
A gas turbine engine defining a radial direction, the gas turbine engine comprising:
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
a fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFanLE and a trailing edge fan radius and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, wherein the leading edge fan radius RFan LE is greater than or equal to 20 inches and less than or equal to 85 inches, the gas turbine engine defining a bypass ratio equal to a mass flowrate of an airflow from the fan over the turbomachine to a mass flowrate of an airflow from the fan through the inlet to the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100;
an outer nacelle at least partially surrounding the fan; and
a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan, the reduction gearbox defining a gear ratio greater than or equal to 2 and less than or equal to 4;
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8, the FLTCF being equal to:
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A gas turbine engine defining a radial direction, the gas turbine engine comprising:
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
a fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFan_LE and a trailing edge fan radius RFan_TE, and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, the gas turbine engine defining a bypass ratio equal to a mass flowrate of an airflow from the fan over the turbomachine to a mass flowrate of an airflow from the fan through the inlet to the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100;
a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan,
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8, the FLTCF being equal to:
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and a gearbox efficiency rating of 0.10-1.8, wherein the gearbox efficiency rating equals QD1.56T1.53 , wherein Q is a gearbox oil flow rate at an inlet of the reduction gearbox measured in gallons per minute at a max takeoff condition, wherein D is a diameter of the fan measured in inches, and wherein T is a net thrust of the gas turbine engine measured in pounds force at the max takeoff condition.
Claim 9
The gas turbine engine of claim 1, wherein the leading edge fan radius RFan_LE is greater than or equal to 35 inches and less than or equal to 50 inches, wherein the fan defines a fan blade count greater than or equal to 12 and less than or equal to 23, and wherein the reduction gearbox defines a gear ratio between 2:1 and 4:1.
Claim 1 provisionally rejected on the ground of nonstatutory double patenting as being unpatentable over claim 5/14 of copending Application No. 19/545,467 (reference application), as indicated below. Although the claims at issue are not identical, they are not patentably distinct from each other.
This is a provisional nonstatutory double patenting rejection because the patentably indistinct claims have not in fact been patented.
Please see the comparison table below; the underlined recitations indicate a feature not claimed in the other patent document.
Instant claim 1 is rejected relative to claim 5/14
Copending claim 5/14 discloses all of Instant Claim 1 except for a nacelle surrounding the fan, however gas turbines without a fan nacelle (open rotor) and gas turbines with a fan nacelle (closed rotor) are well known alternatives in the art.
Instant Application 19/362,542
Copending Application 19/545,467
Claim 1
Claim 1
A gas turbine engine defining a radial direction, the gas turbine engine comprising:
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
a fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFanLE and a trailing edge fan radius and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, wherein the leading edge fan radius RFan LE is greater than or equal to 20 inches and less than or equal to 85 inches, the gas turbine engine defining a bypass ratio equal to a mass flowrate of an airflow from the fan over the turbomachine to a mass flowrate of an airflow from the fan through the inlet to the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100;
an outer nacelle at least partially surrounding the fan; and
a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan, the reduction gearbox defining a gear ratio greater than or equal to 2 and less than or equal to 4;
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8, the FLTCF being equal to:
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A gas turbine engine defining a radial direction, the gas turbine engine comprising:
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
a fan having a fan blade of a plurality of fan blades formed of a composite material, the fan blade defining a leading edge fan radius RFan_LE and a trailing edge fan radius RFan_TE, and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, the gas turbine engine defining a bypass ratio equal to a mass flowrate of an airflow from the fan over the turbomachine to a mass flowrate of an airflow from the fan through the inlet to the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100; and
a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan;
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8, the FLTCF being equal to:
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wherein the fan defines a blade effective acoustic length (BEAL) defined as: BEAL=2c2S1-rrNbcosγ, wherein c is the chord length of the fan blade, S is a span of the fan blade, rr is a radius ratio of the fan, γ is a stagger angle of the fan blade, and Nb is the number of the plurality of fan blades; a nacelle that includes a fan case that surrounds the fan; a plurality of outlet guide vanes including an outlet guide vane, the plurality of outlet guide vanes disposed aft of the fan and extending radially between the turbomachine and the fan case, wherein the gas turbine engine defines an acoustic spacing from a fan blade trailing edge of the fan blade to an outlet guide vane leading edge of the outlet guide vane, wherein the gas turbine engine further defines an acoustic spacing ratio (ASR) defined as: ASR=1NvNb∙AsBEAL wherein As is the acoustic spacing and Nv is the number of the plurality of outlet guide vanes, and wherein the ASR of the gas turbine engine is 1.5 to 16.0.
Claim 1 continued
Claim 5
(see above)
The gas turbine engine of claim 1, wherein the leading edge fan radius RFan_LE is greater than or equal to 35 inches and less than or equal to 50 inches, wherein the fan defines a fan blade count greater than or equal to 12 and less than or equal to 23, and wherein the reduction gearbox defines a gear ratio between 2:1 and 4:1.
Claim 1 continued
Claim 14
(see above)
The gas turbine engine of claim 1, wherein the leading edge fan radius RFan_LE is greater than or equal to 35 inches and less than or equal to 50 inches, wherein the fan defines a fan blade count greater than or equal to 12 and less than or equal to 22, and wherein the gearbox assembly has a gear ratio that is equal to or greater than 2:1 and equal to or less than 4:1.
Claim 1 provisionally rejected on the ground of nonstatutory double patenting as being unpatentable over claim 4 of copending Application No. 19/558,663 (reference application), as indicated below. Although the claims at issue are not identical, they are not patentably distinct from each other.
This is a provisional nonstatutory double patenting rejection because the patentably indistinct claims have not in fact been patented.
Please see the comparison table below; the underlined recitations indicate a feature not claimed in the other patent document.
Instant claim 1 is rejected relative to claim 4
Copending claim 4 discloses all of Instant Claim 1 except for a nacelle surrounding the fan, however gas turbines without a fan nacelle (open rotor) and gas turbines with a fan nacelle (closed rotor) are well known alternatives in the art.
Instant Application 19/362,542
Copending Application 19/558,663
Claim 1
Claim 4
A gas turbine engine defining a radial direction, the gas turbine engine comprising:
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
a fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFanLE and a trailing edge fan radius and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, wherein the leading edge fan radius RFan LE is greater than or equal to 20 inches and less than or equal to 85 inches, the gas turbine engine defining a bypass ratio equal to a mass flowrate of an airflow from the fan over the turbomachine to a mass flowrate of an airflow from the fan through the inlet to the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100;
an outer nacelle at least partially surrounding the fan; and
a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan, the reduction gearbox defining a gear ratio greater than or equal to 2 and less than or equal to 4;
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8, the FLTCF being equal to:
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A gas turbine engine defining a radial direction, the gas turbine engine comprising:
a fan;
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
a first stage of composite fan blades having a first blade comprising:
a first composite body extending chordwise from a first body leading edge to a first body trailing edge; and
a first leading edge protector having a first protector leading edge different from, and receiving at least a portion of, the first composite body,
wherein a first leading length (FLL) extends chordwise from the first protector leading edge to a first end of the first leading edge protector, and a first chord length (FCL) extends chordwise from the first protector leading edge to the first body trailing edge,
wherein a leading edge fan radius RFan_LE extends radially from a longitudinal centerline to the first protector leading edge at a fan blade tip and a trailing edge fan radius extends radially from a longitudinal centerline to the first body trailing edge at the fan blade tip, a leading edge hub radius extends radially from a longitudinal centerline to the first protector leading edge at a fan blade root, and a trailing edge hub radius extends radially from a longitudinal centerline to the first body trailing edge at a fan blade root, and
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF), wherein FLTCF=
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and the FLTCF is greater than or equal to 1.05 and less than or equal to 1.8 (1.05 S FLTCF1.8); and
a second stage of composite airfoils downstream of the first stage of composite airfoils and having a second airfoil comprising:a second composite body extending chordwise from a second body leading edge to a second body trailing edge; and a second leading edge protector having a second protector leading edge different from, and receiving at least a portion of, the second composite body,wherein a second leading length (SLL) extends chordwise from the second protector leading edge to a second end of the second leading edge protector, and a second chord length (SCL) extends chordwise from the second protector leading edge to the second body trailing edge;wherein the FLL and the FCL are related to the SLL and the SCL by a stage protection factor (SPF), wherein SPF =and the SPF is greater than or equal to 0.7 and less than or equal to 4 (0.7 s SPF < 4);
wherein the gas turbine engine defines a bypass ratio equal to a mass flowrate of an airflow from the fan over the turbomachine to a mass flowrate of an airflow from the fan through the inlet defining at least a portion of the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100 (10 : bypass ratio 5 100).
Claim 3
The gas turbine engine of claim 1, further comprising a speed reduction device driven by a turbine section of the gas turbine engine and configured for rotating the fan.
Claim 4
The gas turbine engine of claim 3, wherein the leading edge fan radius is greater than or equal to 35 inches and less than or equal to 50 inches, wherein the fan defines a fan blade count greater than or equal to 12 and less than or equal to 23, and wherein the speed reduction device defines a gear ratio greater than or equal to 2:1 (2:1 : gear ratio).
Claim 1 provisionally rejected on the ground of nonstatutory double patenting as being unpatentable over claim 12 of copending Application No. 19/558,747 (reference application), as indicated below. Although the claims at issue are not identical, they are not patentably distinct from each other.
This is a provisional nonstatutory double patenting rejection because the patentably indistinct claims have not in fact been patented.
Please see the comparison table below; the underlined recitations indicate a feature not claimed in the other patent document.
Instant claim 1 is rejected relative to copending claim 12
Copending claim 12 discloses all of Instant Claim 1 except for a nacelle surrounding the fan, however gas turbines without a fan nacelle (open rotor) and gas turbines with a fan nacelle (closed rotor) are well known alternatives in the art.
Instant Application 19/362,542
Copending Application 19/558,747
Claim 1
Claim 12
A gas turbine engine defining a radial direction, the gas turbine engine comprising:
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
a fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFanLE and a trailing edge fan radius and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, wherein the leading edge fan radius RFan LE is greater than or equal to 20 inches and less than or equal to 85 inches, the gas turbine engine defining a bypass ratio equal to a mass flowrate of an airflow from the fan over the turbomachine to a mass flowrate of an airflow from the fan through the inlet to the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100;
an outer nacelle at least partially surrounding the fan; and
a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan, the reduction gearbox defining a gear ratio greater than or equal to 2 and less than or equal to 4;
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8, the FLTCF being equal to:
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A gas turbine engine defining a radial direction, the gas turbine engine comprising:
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
a fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFan_LE and a trailing edge fan radius RFan_TE, and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, the gas turbine engine defining a bypass ratio equal to a mass flowrate of an airflow from the fan over the turbomachine to a mass flowrate of an airflow from the fan through the inlet to the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100; and
a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan;
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8, the FLTCF being equal to:
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wherein each rotating stage of the drive turbine comprises an annular exit area defined by a tip radius of a trailing edge of any one blade of the rotating stage and a hub radius of the any one blade of the rotating stage at an axial location aligned with the tip radius, wherein the drive turbine comprises an area ratio equal to the annular exit area of an aft-most rotating stage of the drive turbine divided by the annular exit area of a forward-most rotating stage of the drive turbine, and wherein the area ratio is within a range of 2.0-6.5.
Claim 12
The gas turbine engine of claim 1, wherein the leading edge fan radius RFan_LE is greater than or equal to 35 inches and less than or equal to 50 inches, wherein the fan defines a fan blade count greater than or equal to 12 and less than or equal to 23, and wherein the reduction gearbox defines a gear ratio between 2:1 and 4:1.
Claim 1, 12 provisionally rejected on the ground of nonstatutory double patenting as being unpatentable over claim 9, 18 of copending Application No. 19/572,969 (reference application), as indicated below. Although the claims at issue are not identical, they are not patentably distinct from each other.
This is a provisional nonstatutory double patenting rejection because the patentably indistinct claims have not in fact been patented.
Please see the comparison table below; the underlined recitations indicate a feature not claimed in the other patent document.
Instant claim 1 is rejected relative to copending claim 9
Copending claim 9 discloses all of Instant Claim 1 except for a nacelle surrounding the fan, however gas turbines without a fan nacelle (open rotor) and gas turbines with a fan nacelle (closed rotor) are well known alternatives in the art.
Instant claim 12 is rejected relative to copending claims 18
Copending claim 18 discloses all of Instant Claim 1 except for a nacelle surrounding the fan, however gas turbines without a fan nacelle (open rotor) and gas turbines with a fan nacelle (closed rotor) are well known alternatives in the art.
Instant Application 19/362,542
Copending Application 19/572,969
Claim 1
Claim 9 (1+9)
A gas turbine engine defining a radial direction, the gas turbine engine comprising:
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
a fan having a fan blade formed of a composite material,
the fan blade defining a leading edge fan radius RFanLE and a trailing edge fan radius and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, wherein the leading edge fan radius RFan LE is greater than or equal to 20 inches and less than or equal to 85 inches, the gas turbine engine defining a bypass ratio equal to a mass flowrate of an airflow from the fan over the turbomachine to a mass flowrate of an airflow from the fan through the inlet to the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100;
an outer nacelle at least partially surrounding the fan; and
a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan, the reduction gearbox defining a gear ratio greater than or equal to 2 and less than or equal to 4;
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8, the FLTCF being equal to:
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A gas turbine engine defining a radial direction, the gas turbine engine comprising:
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
a fan with a plurality of fan blades with a solidity less than 1.6 and greater than 1.0, wherein the solidity is defined by a ratio of an airfoil chord length to a circumferential pitch of the fan blades, wherein the airfoil chord length is measured at 60% of a radial distance from an axial centerline of the fan to a tip of a fan blade of the plurality of fan blades, wherein a fan blade, of the plurality of fan blades, is formed of a composite material,
the fan blade defining a leading edge fan radius RFan_LE and a trailing edge fan radius RFan_TE, and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, the gas turbine engine defining a bypass ratio equal to a mass flowrate of an airflow from the fan over the turbomachine to a mass flowrate of an airflow from the fan through the inlet to the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100; and
a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan;
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8, the FLTCF being equal to:
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Claim 9
The gas turbine engine of claim 1, wherein the leading edge fan radius RFan_LE is greater than or equal to 35 inches and less than or equal to 50 inches, wherein the fan defines a fan blade count greater than or equal to 12 and less than or equal to 23, and wherein the reduction gearbox defines a gear ratio between 2:1 and 4:1.
Claim 12
Claim 18 (12+18)
A gas turbine engine defining a radial direction, the gas turbine engine comprising:
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
an unducted fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFanLE and a trailing edge fan radius RFanTE, and the fan defining a leading edge hub radius RHUb_LE and a trailing edge hub radius RHUb_TE, the gas turbine engine defining a bypass ratio equal to a mass flowrate of an airflow from the unducted fan over the turbomachine to a mass flowrate of an airflow from the unducted fan through the inlet to the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100; and
a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the unducted fan, wherein the reduction gearbox defines a gear ratio greater than 2 and less than 14;
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Opening Ratio (FLTOR) greater than or equal to 1.03 and less than or equal to 1.5, the FLTOR being equal to:
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A gas turbine engine defining a radial direction, the gas turbine engine comprising:
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
a fan with a plurality of fan blades with a solidity less than 1.6 and greater than 1.0, wherein the solidity is defined by a ratio of an airfoil chord length to a circumferential pitch of the fan blades, wherein the airfoil chord length is measured at 60% of a radial distance from an axial centerline of the fan to a tip of a fan blade of the plurality of fan blades; wherein the fan blade is formed of a composite material, the fan blade defining a leading edge fan radius RFan_LE and a trailing edge fan radius RFan_TE, and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, the gas turbine engine defining a bypass ratio equal to a mass flowrate of an airflow from the fan over the turbomachine to a mass flowrate of an airflow from the fan through the inlet to the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100; and
a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan;
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Opening Ratio (FLTOR) greater than or equal to 1.03 and less than or equal to 1.5, the FLTOR being equal to:
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Claim 18
The gas turbine engine of claim 12, wherein the leading edge fan radius RFan_LE is greater than or equal to 35 inches and less than or equal to 50 inches, wherein the fan defines a fan blade count greater than or equal to 12 and less than or equal to 23, wherein the reduction gearbox defines a gear ratio between 2:1 and 4:1, and wherein the FLTOR is greater than or equal to 1.07 and less than or equal to 1.18.
Allowable Subject Matter
The following is a statement of reasons for the indication of allowable subject matter over art, the office notes that there are pending Double Patenting rejections detailed above:
Claim 1
The prior art of record fails to anticipate or render obvious the limitations of the claim, and in particular the limitations cited below in combination with the remaining limitations of the claim.
“
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”
Claim 12
The prior art of record fails to anticipate or render obvious the limitations of the claim, and in particular the limitations cited below in combination with the remaining limitations of the claim.
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”
Claims 2-4, 9-11, 13-14, 17, 19-24 are allowable over art based on dependency.
Conclusion
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/JOHN S HUNTER, JR/Examiner, Art Unit 3761
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