Petition for Inter Partes Review of U.S. 9,121,412
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`UNITED STATES PATENT AND TRADEMARK OFFICE
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`———————
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`BEFORE THE PATENT TRIAL AND APPEAL BOARD
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`———————
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`General Electric Company,
`Petitioner,
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`v.
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`United Technologies Corporation,
`Patent Owner
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`———————
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`
`
`PETITION FOR INTER PARTES REVIEW
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`OF
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`U.S. PATENT NO. 9,121,412
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`Petition for Inter Partes Review of U.S. 9,121,412
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`Table of Contents
`I.  MANDATORY NOTICES ............................................................................. 1 
`A. 
`Real Party-in-Interest ............................................................................ 1 
`B. 
`Related Matters ...................................................................................... 1 
`C. 
`Lead and Back-up Counsel and Service Information ........................... 1 
`II.  GROUNDS FOR STANDING ........................................................................ 1 
`III. 
`INTRODUCTION ........................................................................................... 2 
`A.  Overview of Gas Turbine Engine Technology ..................................... 4 
`1. 
`Engine Architecture .................................................................... 4 
`2. 
`Fan Section .................................................................................. 7 
`The 412 Patent ..................................................................................... 10 
`B. 
`IV.  STATUTORY GROUNDS FOR THE CHALLENGES .............................. 13 
`V. 
`CLAIM CONSTRUCTION .......................................................................... 15 
`A. 
`“spool” (claim 1) ................................................................................. 15 
`B. 
`“propulsor” (claims 1 and 9) ............................................................... 16 
`C. 
`“pressure ratio that is between 1.1 and 1.35 with regard to an
`inlet pressure and an outlet pressure of said bypass flow
`passage” (claim 1) ............................................................................... 16 
`D.  Operating Condition of the Claimed Operating Parameters ............... 20 
`IDENTIFICATION OF HOW THE CLAIMS ARE
`UNPATENTABLE ........................................................................................ 20 
`A.  Ground 1: Claims 1, 2, 4, 5 and 7-10 are Anticipated by Davies ...... 20 
`1. 
`Claim 1 ...................................................................................... 22 
`2. 
`Claims 2 and 4 ........................................................................... 30 
`3. 
`Claim 5 ...................................................................................... 31 
`4. 
`Claims 7-10 ............................................................................... 32 
`B.  Ground 2: Claims 1, 2, 4, 5, 7, 8, and 11 are rendered obvious
`by Davies in view of the Knowledge of One of Ordinary Skill
`in the Art .............................................................................................. 34 
`1. 
`Claims 1, 2, 4, 5, 7, and 8 ......................................................... 35 
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`VI. 
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`Petition for Inter Partes Review of U.S. 9,121,412
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`Claim 11 .................................................................................... 39 
`2. 
`C.  Ground 3: Claim 5 is Rendered Obvious Based on Davies in
`View of Middleton .............................................................................. 41 
`D.  Ground 4: Claims 1, 3, and 4 are Rendered Obvious by
`Schaefer in view of the Knowledge of One of Ordinary Skill in
`the Art .................................................................................................. 43 
`1. 
`Claim 1 ...................................................................................... 45 
`2. 
`Claim 3 ...................................................................................... 53 
`3. 
`Claim 4 ...................................................................................... 54 
`VII.  CONCLUSION .............................................................................................. 55 
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`Petition for Inter Partes Review of U.S. 9,121,412
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`PETITIONER’S EXHIBIT LIST
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`GE-1006
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`GE-1001 U.S. Patent No. 9,121,412
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`GE-1002 Prosecution File History of U.S. Patent No. 9,121,412
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`GE-1003 Declaration of Reza Abhari Under 37 C.F.R. § 1.68.
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`GE-1004 Curriculum Vitae of Reza Abhari
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`GE-1005 D.G.M. Davies, et al., A Variable Pitch Fan for an Ultra Quiet
`Demonstrator Engine (1976)
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`614: VFW’s Jet Feedliner, Flight International (November 4, 1971)
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`GE-1007 U.S. Patent No. 7,374,403 to Decker, et al.
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`GE-1008 NASA SP-7037 (92), A Cumulative Index to the 1977 Issues of
`Aeronautical Engineering: A Special Bibliography (January 1978)
`(excerpt)
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`John W. Schaefer et al., Dynamics of High-Bypass-Engine Thrust
`Reversal Using A Variable-Pitch Fan (May 1977).
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`GE-1010 NASA Technical Reports Server Record Details for GE-1016
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`GE-1011 William S. Willis, Quiet Clean Short-Haul Experimental Engine
`(QCSEE) Final Report (August 1979).
`
`GE-1012 Bill Sweetman et al., Pratt & Whitney’s surprise leap, INTERAVIA
`(June 1998).
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`GE-1013 Gerald Brines, The Turbofan of Tomorrow, Mechanical Engineering
`(August 1990).
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`GE-1014 Excerpts from Jack D. Mattingly, Elements of Gas Turbine
`Propulsion (1996).
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`GE-1009
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`Petition for Inter Partes Review of U.S. 9,121,412
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`GE-1015 Bill Gunston, Pratt & Whitney PW8000, Jane’s Aero-Engines Issue 7
`(March 2000).
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`GE-1016 Bruce E. Wendus et al., Follow-On Technology Requirement Study
`for Advanced Subsonic Transport (August 2003).
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`GE-1017 Richard Whitaker, ALF502: plugging the turbofan gap, Flight
`International (Jan. 30, 1982).
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`GE-1018 About NASA Technical Reports Server (www.sti.nasa.gov/find-sti).
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`GE-1019 University of California at Davis MARC record for Davies
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`GE-1020 NASA Technical Reports Server Record Details for Schaefer
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`GE-1021 U.S. 5,141,400 to Murphy et al.
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`GE-1022 S.A. Savelle et al., Application of Transient and Dynamic
`Simulations to the U.S. Army T55-L-712 Helicopter Engine (1996).
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`GE-1023 A Summary of Commonly Used Marc 21 Authority Fields, Library of
`Congress
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`Petition for Inter Partes Review of U.S. 9,121,412
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`I. MANDATORY NOTICES
`A. Real Party-in-Interest
`The real parties-in-interest are General Electric Company (Petitioner),
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`Snecma S.A., CFM International S.A., CFM International, Inc., and Safran S.A.
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`B. Related Matters
`As of the filing of this Petition and to the best knowledge of the Petitioner,
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`U.S. Patent No. 9,121,412 (the “412 Patent”) is not involved in litigation.
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`C. Lead and Back-up Counsel and Service Information
`Petitioner designates Anish Desai (Reg. No. 73,760), available at 1300 Eye
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`Street NW, Suite 900, Washington, DC 20005 (T: 202-682-7103) as lead counsel,
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`and Brian Ferguson (Reg. No. 36,801) (T: 202-682-7516) and Christopher Pepe
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`(Reg. No. 73,851) (T: 202-682-7153), available at the same address, as back-up
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`counsel. Please address all correspondence to both lead and back-up counsel.
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`Petitioner consents to service by electronic email at the following address:
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`GE.412.IPR@weil.com.
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`II. GROUNDS FOR STANDING
`Petitioner certifies that the 412 Patent is available for inter partes review and
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`that Petitioner is not barred or estopped from requesting inter partes review
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`challenging the patent claims on the grounds identified in this Petition.
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`III.
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`Petition for Inter Partes Review of U.S. 9,121,412
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`INTRODUCTION
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`The 412 Patent claims a conventional geared turbofan engine configuration
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`having a turbine coupled to drive a shaft, a fan driven by the shaft, and a gear
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`assembly situated between the fan and the shaft to allow the fan to rotate slower
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`than the shaft. This type of turbofan configuration has been known in the art for
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`decades.
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`For example, a 1976 publication by Rolls Royce (“Davies”) (GE-1005)
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`describes the M45SD-02 demonstrator geared turbofan engine. As shown in the
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`figure below, this geared engine included a turbine to drive a shaft (“Low Spool”
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`below), a fan driven by the shaft, and a gear assembly between the fan and the
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`shaft to allow the fan to rotate slower than the shaft.
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`GE-1005.019, Figure 1(annotations in color); GE-1003 at ¶ 71
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`In addition to the conventional elements of a geared turbofan engine, the 412
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`Patent requires a specific configuration of the fan: no more than 16 fan blades, a
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`solidity (R)1 between 0.6 and 0.9, and a ratio of N/R, where N is the number of
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`Petition for Inter Partes Review of U.S. 9,121,412
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`blades, between 8 and 16 or between 18 and 28. The claims also require a pressure
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`ratio, between 1.1 and 1.35, with regard to the inlet and outlet of the bypass flow
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`passage. As explained in detail below, Davies discloses a fan with 14 blades, a
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`solidity (R) equal to 0.74, and a ratio of N/R equal to 18.9. Davies further
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`discloses a fan pressure ratio between 1.18 and 1.27. Similarly, a 1977 NASA
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`publication (“Schaefer”) (GE-1009) discloses a gear-driven fan engine having a
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`fan with 13 blades, a solidity (R) equal to 0.67, a ratio of N/R equal to 19.4, and a
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`fan pressure ratio of 1.18.
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`Neither Davies nor Schaefer was of record during prosecution of the 412
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`Patent. Davies anticipates and/or renders obvious claims 1, 2, 4, 5 and 7-11, while
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`Schaefer renders obvious claims 1, 3, and 4. Petitioner accordingly requests that
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`the Board institute inter partes review of claims 1-5 and 7-11, and cancel these
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`claims as invalid. The remainder of this Petition describes the 412 Patent and its
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`claims, the prior art cited in the Petition, and the reasons why claims 1-5 and 7-11
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`are invalid in light of the prior art.
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`1 Solidity is defined in the claims as the chord dimension at the tip of the blade
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`(CD) divided by the circumferential pitch at the tip of the blades (CP). GE-1001 at
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`4:55-60.
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`Petition for Inter Partes Review of U.S. 9,121,412
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`A. Overview of Gas Turbine Engine Technology2
`1.
`Engine Architecture
`The 412 Patent describes a turbofan engine that is generally used for
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`powering airplanes. See e.g., GE-1001 at 1:15-18 (“This disclosure relates to gas
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`turbine engines and, more particularly, to an engine having a geared turbo fan
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`architecture”); GE-1014.024 (“turbofan engines used in subsonic aircraft”). As
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`disclosed in the prior art, a turbofan engine may include a fan, compressor section,
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`combustion section, and turbine section. GE-1003 [Abhari Decl.] at ¶ 21. The
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`figure below from a 1996 textbook illustrates a conventional two-spool turbofan
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`configuration. Id. at ¶¶ 21, 28.
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`GE-1014.024, Figure 1-7 (annotations in color); GE-1003 at ¶ 21
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`2 Dr. Abhari’s expert declaration includes a more in-depth overview of gas turbine
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`technology. GE-1003 at ¶¶ 18-42.
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`Petition for Inter Partes Review of U.S. 9,121,412
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`As shown above, air enters through the fan section, which is configured to
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`deliver a portion of air into the core flow path (orange), and a portion of air into a
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`bypass duct (blue). GE-1003 at ¶ 22. The air that enters the bypass duct bypasses
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`the core engine to generate thrust. Id. at ¶ 24. The core flow, on the other hand,
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`travels through the compressor section, combustor section, and turbine section
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`before exiting the engine via the exhaust section. Id. It is standard practice to refer
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`to the ratio of the mass flow rate of air bypassing the engine core (i.e., bypass flow)
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`to the mass flow rate of air passing through the engine core (i.e., core flow) as the
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`bypass ratio. Id. at ¶ 22.
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`The core flow that enters the compressor section is compressed by a low
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`pressure compressor and a high pressure compressor. Id. at ¶ 24. The compressed
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`air is subsequently mixed with fuel and ignited in the combustor section. Id. at ¶
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`25. After exiting the combustor section, the core flow is expanded through the
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`high pressure turbine, which drives the high pressure compressor via the high spool
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`shaft. Id. at ¶ 26. The flow is further expanded through the low pressure turbine,
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`which drives the low pressure compressor and the fan via the low spool shaft. Id.
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`at ¶ 27.
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`In a “direct drive” turbofan engine, the low pressure turbine drives the low
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`pressure compressor and fan via a low spool shaft. Id. at ¶¶ 29-30. The low spool
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`components—low pressure turbine, low pressure compressor and fan—all rotate at
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`5
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`the same speed. Id. A geared turbofan differs from a direct drive turbofan because
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`Petition for Inter Partes Review of U.S. 9,121,412
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`it incorporates a speed reduction mechanism (i.e., a gearbox) between the fan
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`section and the components of the low spool. Id. at ¶ 31. The gearbox is
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`connected to the fan section on one side, and the low spool shaft on the other side.
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`Id. The low pressure turbine drives the fan section through the gearbox, which
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`enables the fan to rotate at a lower rotational speed than the low pressure turbine
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`and low pressure compressor. GE-1012.002 (“Gearing solves the problem. The LP
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`turbine and compressor spin faster…. The fan, gear driven off the LP spool, will
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`rotate at about one-third compressor speed….”); see also GE-1003 at ¶ 31. A
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`geared turbofan configuration is illustrated in the figure below from a 1990
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`publication:
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`GE-1013.006 (annotations in red); GE-1003 at ¶ 33
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`It has been well understood in the aviation industry for decades that a geared
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`turbofan configuration can offer certain benefits relative to a direct drive
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`configuration. GE-1003 at ¶¶ 34-35. For example, a geared configuration can
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`enable a high bypass ratio, which can yield improvements in noise levels and fuel
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`efficiency. Id. It is also well known that high bypass ratio engines are typically
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`optimized at a fan pressure ratio that is low (as compared with the fan pressure
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`ratio of low bypass ratio engines). Id.
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`Fan Section
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`2.
`A conventional turbofan engine includes a fan having a number of blades
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`circumferentially spaced around a hub. Id. at ¶ 37. As shown in the figure below,
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`each blade extends radially outward from the hub between a root and a tip, and
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`extends in a chord direction between a leading edge and a trailing edge. Id. The
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`chord dimension can vary over the span of the blade. In other words, the chord
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`dimension at the root (CDroot) is different than the chord dimension at the tip
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`(CDtip). Id.
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`GE-1003 at ¶ 37
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`The circumferential pitch of the blades generally refers to the spacing
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`between the blades. Id. at ¶ 38. If the number of blades and the diameter of the
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`fan are known, then the circumferential pitch (CP) of the blades at the tip can be
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`determined by dividing the circumference of the fan by the number of blades:
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`CP(cid:3404) π ∗ Fan Diameter
`Number of Blades
`chord dimension (cid:4666)CD(cid:4667)
`Solidity (cid:4666)R(cid:4667)(cid:3404)
`circumferential pitch (cid:4666)CP(cid:4667)
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`Id. The ratio of the chord dimension (CD) to the circumferential pitch (CP) is
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`known in the art as solidity.
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`Id. The 412 Patent describes and claims a solidity value as measured at the tip of
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`the fan blade.3 GE-1001 at 4:56-59 (“extends…between a leading edge and a
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`trailing edge at the tip to define a chord dimension (CD), said row of propulsor
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`blades defining a circumferential pitch (CP) with regard to said tips”).
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`Solidity is a well-known design parameter that characterizes how much area
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`the blades sweep through. Id. at ¶ 39. Lower solidity means less area is swept by
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`the fan blades, while higher solidity means more area is swept by the fan blades.
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`Id. Adjusting the solidity of the fan can have various effects. Increasing solidity
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`by increasing chord dimension, for example, can improve the stability and
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`efficiency of the fan. Id. at ¶ 40. Increasing the chord dimension, however, also
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`increases the size and weight of the fan. Id. Conversely, decreasing the solidity of
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`the fan by either reducing the chord dimension or number of blades reduces the
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`size and weight of the fan, but can also reduce the efficiency and stability of the
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`fan. Id.
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`The 412 Patent describes and claims the ratio of the number of blades to the
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`tip solidity (N/R). As explained by Dr. Abhari, the ratio of N/R is equivalent to the
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`ratio of the circumference of the fan to the chord dimension at the tip. Id. at ¶ 41.
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`3 A person of ordinary skill in the art would understand that the solidity of the fan
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`section varies depending on the radial location (e.g., root or tip) at which the chord
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`dimension and circumferential pitch are measured. GE-1003 at ¶ 38.
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`The 412 Patent explains that the ratio of N/R can be between 8 and 28. GE-1001
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`at 3:37-40. Put differently, this means that the fan circumference is between 8 to
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`28 times the blade chord dimension at the tip. GE-1003 at ¶ 41. It was well-
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`known in the art to design a fan such that its circumference is between 8 and 28
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`times the blade chord dimension. Id. at ¶ 42. For example, U.S. Patent No.
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`5,141,400, which issued in 1992, states in the Background of the Invention that
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`“[t]he next generation of commercial high thrust engines will have fan diameters
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`ranging in size from 106 inches to 124 inches” and that “[t]ypical fan blades
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`currently have tip chords of about 8 to 12 inches, while the wide chord fan blades
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`for the larger engines will have tip chords in the range of about 20 to 28 inches.”
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`GE-1021 at 1:19-29; GE-1003 at ¶ 42. A fan diameter of 106 inches with a blade
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`chord of 20 inches means the fan circumference is approximately 16 times the
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`blade chord (i.e., N/R is approximately 16). GE-1003 at ¶ 42. A fan diameter of
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`124 inches with a blade chord of 28 inches means the fan circumference is
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`approximately 13 times the blade chord (i.e., N/R is approximately 13). Id.
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`The 412 Patent
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`B.
`The 412 Patent issued from U.S. Patent Application No. 13/176,255 (“255
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`Application”), which was filed on July 5, 2011, and does not claim priority to any
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`other patent applications. The Notice of Allowance included an Examiner’s
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`Amendment, which amended claim 1 to include the limitations that are underlined
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`below:4
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`Petition for Inter Partes Review of U.S. 9,121,412
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`1. A gas turbine engine comprising:
`a spool;
`a turbine coupled to drive the spool;
`a propulsor coupled to be driven by said turbine through said spool;
`and
`a gear assembly coupled between said propulsor and said spool such
`that rotation of said spool drives said propulsor at a different speed
`than said spool,
`wherein said propulsor includes a hub and a row of propulsor blades
`that extend from said hub, and said row includes a number (N) of
`said propulsor blades that is no more than 16, and the propulsor is
`located at an inlet of a bypass flow passage having a pressure ratio
`that is between 1.1 and 1.35 with regard to an inlet pressure and an
`outlet pressure of said bypass flow passage;
`wherein each of said propulsor blades extends radially between a root
`and a tip and in a chord direction between a leading edge and
`trailing edge at the tip to define a chord dimension (CD), said row
`of propulsor blades defining a circumferential pitch (CP) with
`regard to said tips, wherein said row of propulsor blades has a
`solidity value (R) defined as CD/CP that is between 0.6 and 0.9,
`and a ratio of N/R is between 8 and 16 or between 18 and 28.
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`GE-1002.012-.013. The Examiner provided the following reasons for allowance:
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`4 A similar amendment was also made to application claim 15 (which is issued
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`claim 9).
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`“The closest prior art, Decker et al. (US Pat. 7,374,403), neither teaches or makes
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`Petition for Inter Partes Review of U.S. 9,121,412
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`obvious the following in combination with the other claim limitations: ‘a row of
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`propulsor blades defining a circumferential pitch (CP) with regard to said tips,
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`wherein said row of propulsor blades has a solidity value (R) defined as CD/CP
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`that is between 0.6 and 0.9, and a ratio of N/R is between 8 and 16 or between 18
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`and 28.’” Id. at .014-.015. Decker discloses a turbofan engine having either 18 or
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`20 fan blades, and a solidity value between 1.0 and 1.2. GE-1007 at 9:9-18.
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`Dependent claim 2 of the 412 Patent requires that the claimed pressure ratio
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`is between 1.2 and 1.3, while dependent claim 4 requires a pressure ratio between
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`1.1 and 1.2.
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`Dependent claim 3 requires a bypass ratio of approximately 18.
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`Dependent claim 4 requires a design pressure ratio between 1.1 and 1.2.
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`Dependent claim 5 claims the conventional components of a two-spool
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`engine: a low pressure compressor and low pressure turbine coupled to a spool;
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`and a high pressure compressor and high pressure turbine coupled to another spool.
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`Dependent claim 7 requires a number of propulsor blades between 10 and
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`16, while dependent claim 8 requires that the number of propulsor blades is even.
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`Independent claim 9 includes the same elements with respect to the
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`propulsor as claim 1, but does not include limitations for a spool, turbine, gear
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`assembly or pressure ratio.
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`Petition for Inter Partes Review of U.S. 9,121,412
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`Dependent claim 10 requires that the ratio of N/R is between 12 and 16 or
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`between 18 and 20, while dependent claim 11 requires that the ratio of N/R is
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`between 15 and 16.
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`IV. STATUTORY GROUNDS FOR THE CHALLENGES
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`This Petition provides the following challenges:
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`Ground 1 Anticipation of claims 1, 2, 4, 5 and 7-10 under 35 U.S.C. § 102(b)
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`based on D.G.M. Davies et al., A Variable Pitch Fan for an Ultra
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`Quiet Demonstrator Engine (“Davies”) (GE-1005).
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`Ground 2 Obviousness of claims 1, 2, 4, 5, 7, 8, and 11 under 35 U.S.C. § 103
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`based on Davies in view of the knowledge of one of ordinary skill in
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`the art.
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`Ground 3 Obviousness of claim 5 under 35 U.S.C. § 103 based on Davies in
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`view of P. Middleton, 614: VFW’s jet feederline (“Middleton”) (GE-
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`1006).
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`Ground 4 Obviousness of claims 1, 3, and 4 under 35 U.S.C. § 103 based on
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`John W. Schaefer et al., Dynamics of High-Bypass-Engine Thrust
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`Reversal Using A Variable-Pitch Fan (“Schaefer”) (GE-1009) in view
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`of the knowledge of one of ordinary skill in the art.
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`Davies (GE-1005) is an article excerpted from the Proceedings of the “Seeds
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`for Success in Civil Aircraft Design in the Next Two Decades” Convention that
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`was held in May 1976. GE-1005.001. The copy of Davies included as GE-1005
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`was obtained from the University of California at Davis (“UCD”) library. Id.
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`Exhibit GE-1019 is the UCD MARC (Machine-Readable Cataloging) record
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`indicating that Davies has been catalogued at the UCD library since May 12, 2002.
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`GE-1019.001 (Field “008” with a date of “020512”); GE-1023.011-.012
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`(explaining that the first six characters (i.e., 00-05) in Field 008 of a MARC record
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`correspond to the date the record was entered on file (i.e., YYMMDD)). Davies is
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`thus prior art under 35 U.S.C. § 102(b).
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`The article 614: VFW’s jet feederline by P. Middleton (GE-1006) was
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`published in the November 4, 1971 issue of FLIGHT International, which is a
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`weekly magazine that has been in publication since 1909. GE-1006.003
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`(“Founded in 1909… First aeronautical weekly in the world”). The Middleton
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`article is thus prior art under 35 U.S.C. § 102(b).
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`Schaefer (GE-1009) is a NASA Technical Memorandum that is dated May
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`1977. GE-1009.002. The NASA Technical Reports Server5 indicates that
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`Schaefer was published on May 1, 1977, and was acquired by the publicly
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`5 The NASA Technical Reports Server is a searchable online database that
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`provides access to full-text documents, images and videos. GE-1018.001. The
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`documents in the database include those created or funded by NASA between 1958
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`and the present. Id.
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`14
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`accessible database on November 18, 1995. GE-1020.001. Schaefer is therefore
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`Petition for Inter Partes Review of U.S. 9,121,412
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`prior art under § 102(b).
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`V. CLAIM CONSTRUCTION
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`This Petition analyzes the claims consistent with the broadest reasonable
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`interpretation in light of the specification. See 37 C.F.R. § 42.100(b).
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`“spool” (claim 1)
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`A.
`The 412 Patent claims a “spool” in a gas turbine engine. The term “spool”
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`
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`can generally refer in the art to an assembly including a turbine, a compressor, and
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`a connecting shaft. GE-1003 at ¶ 56. However, in the context of the claims and
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`specification of the 412 Patent, the term “spool” would be understood by one of
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`skill in the art to refer to the shaft to which the turbine is connected that drives the
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`propulsor via the gearbox (rather than the assembly consisting of the shaft, turbine
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`and compressor). Id. at ¶¶ 56-57.
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`
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`First, the claims separately require a “spool,” a “turbine coupled to drive the
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`spool,” and a “propulsor coupled to be driven by said turbine through said spool.”
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`GE-1001 at 4:40-43. Second, the specification describes a “spool” as including a
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`shaft that couples the compressor and turbine. Id. at 2:28-30 (“The low speed
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`spool 30 generally includes an inner shaft 40 that is coupled with a propulsor 42, a
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`low pressure compressor 44, and a low pressure turbine 46”), 2:35-37 (“The high
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`speed spool 32 includes an outer shaft 50 that is coupled with a high pressure
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`compressor 52 and a high pressure turbine 54.”). Accordingly, one of ordinary
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`Petition for Inter Partes Review of U.S. 9,121,412
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`skill in the art would understand that the term “spool” in claim 1 refers to the
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`connecting shaft. GE-1003 at ¶¶ 56-57.
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`“propulsor” (claims 1 and 9)
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`B.
`Claims 1 and 9 of the 412 Patent require a “propulsor” in a gas turbine
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`engine. The specification explains that “the propulsor 42, which in this example is
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`a fan, includes a rotor 70 having a row 72 of propulsor blades 74 that extend a [sic]
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`circumferentially around a hub 76.” GE-1001 at 2:66-3:2. Accordingly, one of
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`ordinary skill in the art would understand that a fan having a rotor and fan blades
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`arranged circumferentially around a hub falls within the scope of the term
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`“propulsor” as that term is used in the 412 Patent. GE-1003 at ¶ 58.
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`C.
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`“pressure ratio that is between 1.1 and 1.35 with regard to an inlet
`pressure and an outlet pressure of said bypass flow passage” (claim
`1)
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`The claims require a pressure ratio with regard to an inlet pressure of the
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`
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`bypass flow passage and an outlet pressure of the bypass flow passage (hereinafter
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`referred to as “bypass pressure ratio”). As explained below, a person of ordinary
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`skill in the art would understand that: (1) the claimed bypass pressure ratio
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`encompasses the pressure ratio measured at any radial location in the bypass
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`passage (e.g., root or tip), as well as encompassing an average value for the bypass
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`pressure ratio; and (2) the bypass pressure ratio of a turbofan engine having a
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`16
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`conventional bypass duct is substantially equivalent to the fan pressure ratio. GE-
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`Petition for Inter Partes Review of U.S. 9,121,412
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`1003 at ¶¶ 59-63.
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`
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`The inlet (60) and outlet (62) of the bypass flow passage are shown in Figure
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`1 below. See also GE-1001 at 2:51-55 (“[f]or a given design of the propulsor 42,
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`the inlet 60 and the outlet 62, the engine 20 define a design pressure ratio with
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`regard to an inlet pressure at the inlet 60 and outlet pressure at the outlet 62 of
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`bypass flow passage B.”).
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`GE-1001 at Figure 1 (annotations in color); GE-1003 at ¶ 59
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`The claims and specification of the 412 Patent do not specify a radial location (e.g.,
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`hub, tip, or middle of the bypass passage) at which the inlet and outlet pressure of
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`bypass passage are measured. Accordingly, under the broadest reasonable
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`interpretation the claimed bypass pressure ratio encompasses the pressure ratio
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`Petition for Inter Partes Review of U.S. 9,121,412
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`measured at any radial location in the bypass passage, as well as encompassing an
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`average value for the bypass pressure ratio. GE-1003 at ¶ 60.
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`The bypass pressure ratio is related to another engine operating parameter
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`called the fan pressure ratio, which is an operating parameter that is commonly
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`discussed in the technical publications relating to turbofan engines. Id. at ¶ 61.
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`Fan pressure ratio is typically defined as the ratio of the pressure at the outlet of the
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`fan section to the pressure at the inlet of the fan section. Id. at ¶ 62. The figure
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`below from a 1996 textbook provides station numbering for a turbofan engine.
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`The fan inlet is station 2, the fan exit is station 13, and the fan nozzle exit (i.e.,
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`outlet of the bypass passage) is station 19. Id. The fan pressure ratio is the
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`pressure at station 13 divided by the pressure at station 2 (P13 / P2). Id. In
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`comparison, the bypass pressure ratio is the pressure at station 19 divided by the
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`pressure at station 2 (P19 / P2). Id.
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`18
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`Petition for Inter Partes Review of U.S. 9,121,412
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`GE-1014.069, Figure 5-20
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`
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`One of ordinary skill in the art would also understand that for a turbofan
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`engine having a conventional bypass duct6, such as the one shown above, the total
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`pressure drop between station 13 and station 19 is minimal. GE-1003 at ¶ 63.
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`Accordingly, one of ordinary skill in the art would understand that the total
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`pressure at station 13 (P13) will be substantially equivalent to the total pressure at
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`station 19 (P19). Id. This of course means that the fan pressure ratio (P13 / P2) is
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`substantially equivalent to the bypass pressure ratio (P19 / P2). Id.
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`6 A conventional bypass duct means that there are no structures located between
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`station 13 and station 19 that are intended to modify the total pressure of the flow
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`through the bypass flow passage. GE-1003 at ¶ 77.
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`19
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`Petition for Inter Partes Review of U.S. 9,121,412
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`D. Operating Condition of the Claimed Operating Parameters
`A person of ordinary skill would understand that the claimed pressure ratio
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`does not correspond to a particular operating condition of the engine (i.e., ambient
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`temperature, altitude, flight speed, or power level) because the claims of the 412
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`Patent do not specify a particular condition. GE-1003 at ¶ 64.
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`VI.
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`
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`IDENTIFICATION OF HOW THE CLAIMS ARE UNPATENTABLE
`A. Ground 1: Claims 1, 2, 4, 5 and 7-10 are Anticipated by Davies
`Davies is a publication from 1976 that discloses a variable pitch fan for a
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`turbofan engine. Specifically, Davies provides details regarding the M45SD-02
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`Demonstrator Engine depicted below in Figure 1 from Davies:
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`GE-1005.019, Figure 1
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`
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`The M45SD-02 engine is two-spool geared turbofan engine. GE-1005.004
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`(“…when driven by a two spool engine…”); GE-1003 at ¶ 66. As illustrated in
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`Davies, the M45SD-02 includes a low pressure spool having a low shaft
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`20
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`connecting a low pressure compressor and a low pressure turbine to a fan through a
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`Petition for Inter Partes Review of U.S. 9,121,412
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`reduction gearbox, and also includes a high pressure spool having a high shaft
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`connecting a high pressure compressor and a high pressure turbine. GE-1003 at ¶
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`66.
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`Davies also discloses that the geared turbofan has a reduction ratio of 2.38:1,
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`which means that the low pressure turbine and connecting shaft rotate 2.38 times
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`faster than the fan. Id. at ¶ 67. The result of the geared configuration, as would be
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`expected by one of ordinary skill in the art, is a low fan tip speed (1027 feet per
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`second) and a fan pressure ratio that is low (1.18 at hub and 1.27 at tip). GE-
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`1005.005; see also GE-1003 at ¶ 67.
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`Davies also provides details regarding the