DETAILED ACTION
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.
Claims 1-3, 5-17 are rejected on the ground of nonstatutory double patenting as being unpatentable over claims 1-5 and 7-9 of U.S. Patent No. 11,598,223 in view of Moniz 2007/0084183.
Claim 1 of 11,598,223 recites all of the limitations of claims 1-3, 5-9, 11 of the instant application except that the second compressor includes a plurality of stages. Moniz teaches a geared turbofan engine (10) with a high pressure compressor (26) with a plurality of stages (as shown in figure 1, paragraph [0014]). It would have been obvious to one of ordinary skill in the art prior to the filing date of the invention to use multiple stages for the second compressor of claim 1 of 11,598,223, as taught by Moniz, in order to further increase the pressure of the air passing through the compressor before entering the combustion chamber (paragraph [0014]).
Claim 9 of 11,598,223 recites all of the limitations of claim 10 of the instant application except that the second compressor includes a plurality of stages. Moniz teaches a geared turbofan engine (10) with a high pressure compressor (26) with a plurality of stages (as shown in figure 1, paragraph [0014]). It would have been obvious to one of ordinary skill in the art prior to the filing date of the invention to use multiple stages for the second compressor of claim 1 of 11,598,223, as taught by Moniz, in order to further increase the pressure of the air passing through the compressor before entering the combustion chamber (paragraph [0014]).
Claims 2, 3, 7, 8, 4, and 5 of 11,598,223 recites all of the limitations of claims 12-17 of the instant application respectively, except that the second compressor includes a plurality of stages. Moniz teaches a geared turbofan engine (10) with a high pressure compressor (26) with a plurality of stages (as shown in figure 1, paragraph [0014]). It would have been obvious to one of ordinary skill in the art prior to the filing date of the invention to use multiple stages for the second compressor of claim 1 of 11,598,223, as taught by Moniz, in order to further increase the pressure of the air passing through the compressor before entering the combustion chamber (paragraph [0014]).
Claim 4 is rejected on the ground of nonstatutory double patenting as being unpatentable over claim 1 of U.S. Patent No. 11,598,223 in view of Moniz and Henry 2008/0098718.
Claim 1 of 11,598,223 recites all of the limitations of claim 4 of the instant application except that the second compressor includes a plurality of stages and the first compressor comprises fewer stages than the second turbine. Moniz teaches a geared turbofan engine (10) with a high pressure compressor (26) with a plurality of stages (as shown in figure 1, paragraph [0014]). Henry teaches a low pressure turbine (20) of a geared turbofan engine (10) with more stages than a low pressure compressor (22, where LPC comprises 3 stages and LPT comprises 4 stages as shown in figure 1). It would have been obvious to one of ordinary skill in the art prior to the filing date of the invention to use multiple stages for the second compressor of claim 1 of 11,598,223, as taught by Moniz, and to use more second turbine stages than first compressor stages of the turbine engine of claim 4 of 11,598,223, as taught by Henry, in order to further increase the pressure of the air passing through the compressor before entering the combustion chamber (paragraph [0014] of Moniz), and to increase the amount of power extracted from the turbine and transferred to the fan and low pressure compressor (paragraph [0032] of Henry).
Claim Rejections - 35 USC § 103
The following is a quotation of pre-AIA 35 U.S.C. 103(a) which forms the basis for all obviousness rejections set forth in this Office action:
A patent may not be obtained though the invention is not identically disclosed or described as set forth in section 102, if the differences between the subject matter sought to be patented and the prior art are such that the subject matter as a whole would have been obvious at the time the invention was made to a person having ordinary skill in the art to which said subject matter pertains. Patentability shall not be negatived by the manner in which the invention was made.
Claims 1-8, 10, 11, 15-19, and 21 are rejected under pre-AIA 35 U.S.C. 103(a) as being unpatentable over PW1524G (described in Sabnis “The PW1000G Pure Power New Engine Concept and Its Impact on MRO”, EASA Type-Certificate Data Sheet for PW1500G series engines, and cross-section to scale from “Aviation Investigation Report A14Q0068”, with a scale of 23.1:1 based upon the measurements from the report) in view of Sheridan 2010/0105516.
In regards to Independent Claim 1 and Dependent Claims 2-3, 16-18 and 21, PW1524G teaches a gas turbine engine (PW1524G on page 34 of Sabnis) comprising: a propulsor (fan at inlet end on page 34 of Sabnis) including a plurality of blades (as shown on page 34); a compressor section (LPC and HPC on page 34 of Sabnis) including a first compressor (LPC on page 34 of Sabnis) and a second compressor (HPC on page 34 of Sabnis), and the first compressor including a plurality of stages (3 stages on page 34 of Sabnis), and the second compressor including a plurality of stages (8 stages on page 34 of Sabnis); a turbine section (HPT and LPT on page 34 of Sabnis below) including a first turbine (LPT) and a second turbine (HPT), wherein the first turbine drives the second compressor (LPT drives LPC on page 34 of Sabnis), and the second turbine includes three stages (3 stage LPT on page 34 of Sabnis); an epicyclic gear system (fan drive gear system on page 34 of Sabnis) with a gear reduction (gear reduction of approximately 3:1 for the gearbox driving the fan, pages 34 and 36 of Sabnis), wherein the second turbine drives the propulsor through the epicyclic gear system (3:1 gear reduction, page 36 of Sabnis), wherein the epicyclic gear system includes a sun gear, a plurality of intermediate gears surrounding the sun gear, and a ring gear surrounding the intermediate gears (page 37 of Sabnis shows a cross-section of the fan drive gear system with a sun gear, surrounded by planetary gears, which are in turn surrounded by a ring gear), and wherein the gear system is a planetary gear system including a carrier that supports the intermediate gears (carrier that supports planetary gears, page 34 and 37 of Sabnis, attached to static structure as shown on page 34, where the carrier is the output for the gearbox to the fan, and the ring is fixed as is common in all PW1000G engine gearboxes); a drive shaft interconnecting the carrier and the propulsor (shaft between bearings 1 and 1.5 in figure 10 of Aviation Investigation Report below); wherein the second turbine has a first exit area at a first exit point (593.44 square inches based on measurements in figure 10 of Aviation Investigation Report A14Q0068 below) and is rotatable at a first speed (10600 RPM from EASA Type-Certificate Data Sheet for PW1500G series engines), the second turbine has a second exit area at a second exit point (110.6 square inches based on measurements in figure 10 below) and is rotatable at a second speed, the second speed being greater than twice the first speed (PW1524G has LP spool with maximum speed of 10600 RPM, and HP spool with maximum speed of 24470 RPM, see page 12 of EASA Type-Certificate Data Sheet for PW1500G series engines); and wherein a first performance quantity is defined as the product of the first speed squared and the first exit area (6.67 x 10^10 in2rpm2), a second performance quantity is defined as the product of the second speed squared and the second exit area (6.62 x 10^10 in2rpm2), and a performance ratio of the first performance quantity to the second performance quantity is between 0.8 and 1.5 (6.67/6.62 = 1.008). However, PW1524G does not teach a frame supporting the fan drive shaft, and a flexible support supporting the gear system relative to a static structure, the support transverse stiffness between 50 and 80 percent of the frame transverse stiffness, and the support lateral stiffness between 20 and 50 percent of the frame lateral stiffness. Sheridan teaches a frame (68) supporting a portion of a fan drive shaft (supports fan drive shaft through bearings shown at 70 in figure 3), and a flexible support (74) at least partially supporting the gear system (74 is component of 52, which supports the gear system, paragraph [0013]), the flexible support comprising transverse and lateral stiffnesses (physical limitation of the flexible support). Sheridan further teaches that the stiffnesses of the flexible support can be changed by changing the thickness of the support to control displacement of the gear train (paragraph [0030]). Therefore, the transverse and lateral stiffnesses of the flexible support are recognized in the prior art as result effective variables, In re Antonie, 559 F.2d 618, 195 USPQ 6 (CCPA 1977) MPEP 2144.05(II)(B), and therefore could be optimized through routine experimentation. It would have been obvious to one of ordinary skill in the art before the effective filing date of the invention to modify the engine of PW1524G with the frame and flexible support of Sheridan, to support the output shaft from the gearbox (paragraph [0018]) and because it has been held that "Where the general conditions of a claim are disclosed in the prior art, it is not inventive to discover the optimum or workable ranges by routine experimentation.” In re Aller, 220 F.2d 454, 456, 105 USPQ 233, 235 (CCPA 1955) MPEP 2144.05(II)(A), in this case, support transverse and lateral stiffnesses that are between 50 and 65 percent and 20 and 50 percent of the frame transverse and lateral stiffnesses, respectively.
Multiple references were used to describe the same production engine PW1524G because a single reference did not comprise all of the details of the engine relied upon to reject the claims.
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Page 34 of Sabnis
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Figure 10 of Aviation Investigation Report A14Q0068
Regarding Dependent Claim 4, PW1524G in view of Sheridan teaches the invention as claimed and discussed above, and further teaches that the engine is scalable and the low pressure compressor can comprise fewer stages than the low pressure turbine (Page 34 of Sabnis, scalable core and PW1217G provided as example where the LPC comprises 2 stages and the LPT comprises 3 stages). It would have been obvious to one of ordinary skill in the art prior to the filing date of the invention to reduce the number of stages in the low pressure compressor of PW1524G in view of Sheridan, as taught by PW1524G on page 34 of Sabnis, in order to scale the core of the engine dependent upon the amount of thrust needed (page 34 of Sabnis).
Regarding Dependent Claim 5, PW1524G in view of Sheridan teaches the invention as claimed and discussed above, and PW1524G further teaches the epicyclic gear system is straddle- mounted by first and second bearings (bearings 1.5 and 2 on either side of gearbox in figure 10 of Aviation Investigation Report A14Q0068 above) on opposite sides of the gear reduction relative to an engine longitudinal axis (as shown in figure 10 above), the first bearing supports the drive shaft (bearing 1.5 shown supporting fan shaft in figure 10 above), and the second bearing supports an aft portion of the carrier (bearing 2 supports shaft downstream of gearbox, which includes the carrier in figure 10 above).
Regarding Dependent Claim 6, PW1524G in view of Sheridan teaches the invention as claimed and discussed above, and PW1524G further teaches the performance ratio is between 1.0 and 1.075 (1.008 as rejected in claim 1 above).
Regarding Dependent Claim 7, PW1524G in view of Sheridan teaches the invention as claimed and discussed above, and PW1524G further teaches the first compressor includes three stages (3 stages LPC in figure of page 34 above).
Regarding Dependent Claim 8, PW1524G in view of Sheridan teaches the invention as claimed and discussed above, and PW1524G further teaches the second bearing is axially aligned with the carrier relative to the engine longitudinal axis (bearing 2 axially downstream of gearbox with carrier as shown in figure 10 of Aviation Investigation Report A14Q0068 above).
Regarding Dependent Claim 10, PW1524G in view of Sheridan teaches the invention as claimed and discussed above, and PW1524G further teaches that the first bearing is a thrust bearing (tapered roller bearing in figure 10 of “Aviation Investigation Report A14Q0068”) and that the second bearing is a ball bearing (ball bearing in figure 10 of ” Aviation Investigation Report A14Q0068”).
Regarding Dependent Claim 11, PW1524G in view of Sheridan teaches the invention as claimed and discussed above, and PW1524G further teaches the gas turbine engine includes a first spool and a second spool (PW1524G has LP and HP spools turning at 10600 and 24470 RPM respectively, page 12 of EASA Type-Certificate Data Sheet for PW1500G series engines), wherein the first spool includes a first shaft that interconnects the propulsor, the first compressor and the second turbine (LP spool between LPT, LPC and gearbox for fan on page 34 of Sabnis), wherein the second spool includes a second shaft that interconnects the second compressor and the first turbine (HP spool between HPC and HPT on page 34 of Sabnis), and wherein the first shaft and the second shaft are rotatable about the engine longitudinal axis (rotational speeds of spools as rejected above).
Regarding Dependent Claim 15, PW1524G in view of Sheridan teaches the invention as claimed and discussed above, and PW1524G further teaches the sun gear is mounted to a flexible input attached to the first shaft (input is narrowing of shaft into sun gear on figure 10 of Aviation Investigation Report A14Q0068, where flexible is a relative term, and all shafts have some degree of flexibility).
Regarding Dependent Claim 19, PW1524G in view of Sheridan teaches the invention as claimed and discussed above, and PW1524G further teaches the performance ratio is between 1.0 and 1.075 (1.008 as rejected in claim 1 above).
Claims 12-14 are rejected under 35 U.S.C. 103 as being unpatentable over PW1524G in view of Sheridan as applied to claim 11 above, and further in view of Durocher 2010/0132376.
Regarding Dependent Claims 12-14, PW1524G in view of Sheridan teaches the invention as claimed and discussed above, and PW1524G further teaches a mid-turbine frame (Figure 11 of “Aviation Investigation Report A14Q0068”). However, PW1524G in view of Sheridan does not teach that the mid-turbine frame has an overhung bearing supporting the low pressure shaft and an air turning guide vane. Durocher teaches a mid-turbine frame (28, with airfoil vanes 118) having a bearing supporting the low pressure shaft (bearing 104 supports low pressure shaft 12 in overhung manner as shown in figures 1 and 4). It would have been obvious to one of ordinary skill in the art prior to the filing date of the invention to use an overhung bearing to support the low pressure shaft of PW1524G in view of Sheridan at the mid-turbine frame of PW1524G in view of Sheridan, as taught by Durocher, and to place an air turning guide vane in the mid-turbine frame of PW1524G in view of Sheridan, as taught by Durocher, in order to transfer loads from the shaft through the case of the engine (paragraph [0002]) and to direct the combustion gases from the high to low pressure turbine (paragraph [0047]).
Claim(s) 20 is/are rejected under 35 U.S.C. 103 as being unpatentable over PW1524G in view of Sheridan as applied to claim 1 above, and further in view of Moniz 2007/0084183.
Regarding Dependent Claim 20, PW1524G in view of Sheridan teaches the invention as claimed and discussed above. However, PW1524G in view of Sheridan does not teach that the low pressure compressor and fan turbine at a common speed. Moniz teaches positioning a gearbox (100) such that a fan (52) and low pressure compressor (24) turn at a common speed (paragraph [0016]). It would have been obvious to one of ordinary skill in the art prior to the filing date of the invention to connect the low pressure compressor to the fan of PW1524G in view of Sheridan to turn at the same speed, as taught by Moniz, in order to reduce gear losses by driving the low pressure compressor and fan at the same speed (paragraph [0027]).
Response to Arguments
Applicant's arguments filed 8/3/2026 have been fully considered but they are not persuasive.
Applicant argues that no evidence has been presented to show that figure 10 of the Aviation Investigation Report is to scale. The measurements from figure 10 of the Aviation Investigation Report comports with the published dimensions for the length and fan diameter from the EASA Type certificate.
Applicant argues that the operational parameters for the PW1524G engine from the EASA Type-Certificate Data Sheet were published after the priority date, and cannot be relied upon in the rejection of the claims. The data sheet is relied upon because it discloses details of the PW1524G engine, where the engine itself was disclosed prior to the priority date, as evidenced by Sabnis.
Applicant argues that examiner failed to establish that the relationship between the flexibility of the flexible support and frame of PW1524G in view of Sheridan is result effective. Sheridan teaches that the thickness, and therefore the flexibility, of the flexible support (74) is result effective (“The amount of resistance provided by flexible coupling 74 can be set to vary based on different engine configurations. For example, the stiffness of the curved portions between lateral segment 88 and longitudinal segments 86 and 90 can be varied by changing the thickness of coupling 74 or the amount of curvature. Thus, during normal operating conditions of engine 10, flange 82 does not engage shoulder 96, and flexible coupling 74 allows gear train 30 to be displaced. However, during extreme operating conditions, shoulder 96 engages flange 82 to inhibit further displacement of gear train 30 to avoid excessive wear or damage.” paragraph [0030] of Sheridan). Optimizing the thickness, and therefore the flexibility of the flexible coupling results in optimizing the ratio of that flexibility of the coupling to that of a frame supporting the fan drive shaft. Unlike in In re Antonie, cited by applicant, the benefit of optimizing the degree of flexibility of the flexible support are identical to the benefits of optimizing a ratio between the flexibility of a flexible support vs a frame, which is the displacement of the gear train without resulting in excessive wear and damage, as recited above.
Conclusion
Applicant's amendment necessitated the new ground(s) of rejection presented in this Office action. Accordingly, THIS ACTION IS MADE FINAL. See MPEP § 706.07(a). Applicant is reminded of the extension of time policy as set forth in 37 CFR 1.136(a).
A shortened statutory period for reply to this final action is set to expire THREE MONTHS from the mailing date of this action. In the event a first reply is filed within TWO MONTHS of the mailing date of this final action and the advisory action is not mailed until after the end of the THREE-MONTH shortened statutory period, then the shortened statutory period will expire on the date the advisory action is mailed, and any nonprovisional extension fee (37 CFR 1.17(a)) pursuant to 37 CFR 1.136(a) will be calculated from the mailing date of the advisory action. In no event, however, will the statutory period for reply expire later than SIX MONTHS from the mailing date of this final action.
Any inquiry concerning this communication or earlier communications from the examiner should be directed to STEVEN M SUTHERLAND whose telephone number is (571)270-1902. The examiner can normally be reached M-F 8-5.
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If attempts to reach the examiner by telephone are unsuccessful, the examiner’s supervisor, Arthur Hall can be reached at (571) 270 - 1814. The fax phone number for the organization where this application or proceeding is assigned is 571-273-8300.
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/STEVEN M SUTHERLAND/Primary Examiner, Art Unit 3752
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