`
` IN THE UNITED STATES PATENT AND TRADEMARK OFFICE
`
`BEFORE THE PATENT TRIAL AND APPEAL BOARD
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`
`
`In re U.S. Patent No. 9,121,412
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`
`
`Filed:
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`July 5, 2011
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`Issued:
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`September 1, 2015
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`Inventors: Edward J. Gallagher, Jun Jiang, Becky E. Rose, Jason Elliot, Anthony
`
`
`R. Bifulco
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`Assignee: United Technologies Corporation
`
`Efficient, Low Pressure Ratio Propulsor for Gas Turbine Engines
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`
`
`DECLARATION OF REZA ABHARI, PH.D.
`
`Title:
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`
`
`
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`I, Reza Abhari, make this declaration in connection with the petition for
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`inter partes review submitted by Petitioner for U.S. Patent No. 9,121,412 (“the 412
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`Patent”). All statements made herein of my own knowledge are true, and all
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`statements made herein based on information and belief are believed to be true.
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`Although I am being compensated for my time in preparing this declaration, the
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`opinions articulated herein are my own, and I have no stake in the outcome of this
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`proceeding or any related litigation or administrative proceedings.
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`GE v. UTC
`IPR2016-00952
`GE-1034.001
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`

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`I.
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`INTRODUCTION
`1.
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`I am making this declaration at the request of the General Electric
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`Company in the matter of the Inter Partes Review of U.S. Patent No. 9,121,412
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`(the “412 Patent”). I previously submitted a declaration in this proceeding (GE-
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`1003), which lists my educational background and qualifications.
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`2.
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`In the preparation of this declaration, I have reviewed the relevant
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`portions of the following documents:
`
`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-1005 D.G.M. Davies, et al., A Variable Pitch Fan for an Ultra Quiet
`Demonstrator Engine (1976)
`
`GE-1024 U.S. Patent Application No. 2009/0314881 to Sucui et al. (published
`Dec. 24, 2009).
`GE-1025 U.S. Patent No. 3,898,799 to Pollert et al. (1975).
`GE-1026 W.K. Lord et al., Flow Control Opportunities in Gas Turbine
`Engines (2000).
`GE-1028 U.S. Patent No. 3,820,719 to Clark (1974).
`GE-1029 David A. Sagerser et al., Reverse-Thrust Technology for Variable-
`Pitch Fan Propulsion Systems (1978).
`GE-1030 R.M. Denning, Variable Pitch Ducted Fans for STOL Transport
`Aircraft (1971).
`GE-1031 Deposition Transcript of K. Mathioudakis (April 20, 2017).
`GE-1032 N.A. Cumptsy, Compressor Aerodynamics (2004).
`GE-1033 Gunter Wilfert, Geared Fan, Aero-Engine Design: From State of the
`Art Turbofans Towards Innovative Architectures (March 3-7, 2008).
`UTC-2007 Alan H. Epstein, Aeropropulsion for Commercial Aviation in the
`Twenty-First Century and Research Directions Needed (2014).
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`GE v. UTC
`IPR2016-00952
`GE-1034.002
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`UTC-2015 Declaration of K. Mathioudakis
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`In forming my opinions expressed below, I have considered the
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`3.
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`documents listed above, and my knowledge and experience based upon my work in
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`this area as described in GE-1003 and below.
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`II. OVERVIEW OF THE 412 PATENT
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`4.
`
`I disagree with Dr. Mathioudakis’ assertion that the “’412 patent
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`identifies and claims specific combinations of gas turbofan engine features that
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`lower losses in the bypass flow passage that impact propulsor (fan) efficiency.”
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`UTC-2015, ¶ 17. There is no description in the 412 Patent of how to design the
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`bypass flow passage to minimize pressure losses. The 412 Patent does not
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`describe what structures may be in the bypass flow passage (e.g., struts or guide
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`vanes) or how to design those structures to minimize pressure losses. Figure 1 of
`
`the 412 Patent, for example, does not even depict fan exit guide vanes, which are a
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`structure that one of ordinary skill would understand could cause some pressure
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`loss in the bypass flow passage. Likewise, the 412 Patent does not describe how to
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`design the bypass flow passage geometry (i.e., shape, length, etc.) to minimize
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`pressure losses.
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`5.
`
`The claims of the 412 Patent are also not limited to a bypass flow
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`passage with minimal pressure losses. The only structure in the bypass flow
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`GE v. UTC
`IPR2016-00952
`GE-1034.003
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`passage that causes a pressure rise is the fan. Thus, the bypass flow passage
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`pressure ratio is a combination of the pressure rise caused by the fan and the
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`pressure losses that occur in the bypass flow passage following the fan.1 The
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`claims require a particular bypass flow passage pressure ratio, but do not specify a
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`fan pressure ratio range or a range of losses in the bypass flow passages.
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`Accordingly, the claims cover an engine having a high fan pressure ratio with high
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`bypass flow passage pressure losses, a low fan pressure ratio with low losses, and
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`many combinations in between.
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`6.
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`I have not seen any evidence that the bypass flow passage pressure
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`1 The bypass flow passage pressure ratio will be lower than the fan pressure ratio
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`because it accounts for losses not included in the fan pressure ratio. But as I
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`described in my initial declaration, a person of ordinary skill in the art would
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`understand that it is desirable to minimize losses in the bypass flow passage
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`because it improves engine performance (e.g., fuel efficiency). GE-1003, ¶ 77;
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`GE-1016.055. Minimizing pressure losses in the bypass duct of a turbofan engine
`
`had been a basic design objective for turbofan engines for decades before the 412
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`Patent was filed. The lower the losses in the bypass flow passage, the closer the
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`bypass flow passage pressure ratio will be to the fan pressure ratio.
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`GE v. UTC
`IPR2016-00952
`GE-1034.004
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`ratio range claimed in the 412 Patent produces an unexpected result. Rather, it is
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`my opinion that a person of ordinary skill in the art would have already known
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`prior to the filing of the 412 Patent that a high bypass ratio in combination with a
`
`low fan pressure ratio, and thus low bypass flow passage pressure ratio, results in
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`improved propulsive efficiency. GE-1016.011 (“The ADP’s lower fan pressure
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`ratio (FPR) gives it a propulsive efficiency advantage….”); GE-1033.009.2 I have
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`also not seen any evidence that the particular ranges claimed in the 412 Patent for
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`the bypass flow passage pressure ratio have any criticality (e.g., the ranges are
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`linked with improved efficiency as compared to bypass flow passage pressure
`
`ratios that fall outside the claimed ranges). The 412 Patent simply says that the
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`engine “may be designed with a particular design pressure ratio” and that it “may
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`be between 1.1 and 1.35” or “may be between 1.2 and 1.3.” No correlation is
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`drawn between these specific ranges and improved efficiency.
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`III. FAN TIP SOLIDITY DISCLOSED IN DAVIES
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`7.
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`As I explained in my previous declaration, Davies discloses
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`dimensions of the fan from which a person of ordinary skill in the art can calculate
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`a blade tip solidity of 0.74 for the M45SD-02 engine. GE-1003, ¶¶ 80-82. I
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`disagree with Dr. Mathioudakis (see UTC-2015, ¶¶ 54-55) that a person of
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`2 All emphasis added herein, unless otherwise noted.
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`GE v. UTC
`IPR2016-00952
`GE-1034.005
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`ordinary skill in the art would associate the blade tip solidity of 0.83 disclosed in
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`Davies with the M45SD-02 engine.
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`8.
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`As I described during my deposition, the fan tip solidity values of 0.8
`
`and 0.83 reported in Davies are disclosed in a section describing a general design
`
`philosophy and a scale model engine. UTC-2013 at 137:2-10. Davies explains,
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`for example, that a tip solidity of 0.8 would be chosen for a “design pressure ratio
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`of say 1.27:1….” GE-1005.007. On the following page of the same general
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`design philosophy section, Davies discloses a design pressure ratio of 1.27 with tip
`
`and root pressure ratios of 1.36:1 and 1.18:1, and a tip solidity of 0.83:
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`GE v. UTC
`IPR2016-00952
`GE-1034.006
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`GE-1005.008 (annotations in yellow)
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`
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`A person of ordinary skill would understand from the text of Davies, however, that
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`the design pressure ratio for the M45SD-02 fan is not 1.27. In the section titled
`
`“Engine Definition,” Davies discloses a table specifically labeled as data for the
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`the M45SD-02, including a fan inner (i.e., hub) pressure ratio of 1.18 and fan outer
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`(i.e., tip) pressure ratio of 1.27:
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`GE v. UTC
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`GE-1034.007
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`GE-1005.005 (annotations in yellow)
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`This means that the design pressure ratio of the M45SD-02 is between those two
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`values, and likely between 1.21-1.24.
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`9.
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`As described in Davies, a person of ordinary skill in the art would
`
`understand that efficiency, solidity, and fan pressure ratio are interrelated
`
`parameters. GE-1032.018 (“The rise in solidity and fall in aspect ratio can both be
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`attributed in the main to a rise in chord length. With these trends for aspect ratio
`
`and solidity there is the striking rise in pressure rise per stage….”). A person of
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`ordinary skill in the art would understand that for a given fan tip speed, utilizing a
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`higher fan pressure ratio while maintaining fan efficiency is achieved by increasing
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`the solidity of the fan, while utilizing a lower fan pressure ratio would be achieved
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`by decreasing the solidity of the fan. GE-1032.018 (“The evidence suggests that
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`GE v. UTC
`IPR2016-00952
`GE-1034.008
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`for a good compressor near the design point efficiency tends to be slightly lower if
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`the solidity is on the high side…but the pressure rise and operating range are
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`greater.”); GE-1030.003 (“The low solidity of the fan system which limits the
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`design pressure ratio”).
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`10. Because Davies correlates a design pressure ratio of 1.27 with a
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`solidity of 0.8 (or 0.83), a person of ordinary skill in the art would expect a solidity
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`value of less than 0.8 for a design pressure ratio of less than 1.27. This is
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`consistent with the tip solidity of 0.74 based on the M45SD-02 fan dimensions
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`disclosed in Davies.
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`IV. MODIFYING THE TIP CHORD DIMENSION OF A FAN BLADE
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`11. According to Dr. Mathioudakis, a person of ordinary skill would not
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`have been motivated to modify the tip chord dimension of the fan blade disclosed
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`in Davies. UTC-2015, ¶¶ 72-76. As described below, I disagree with Dr.
`
`Mathioudakis.
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`12. Dr. Mathioudakis contends that the fan tip chord dimension (i.e., tip
`
`solidity) could not be modified without a corresponding increase to the fan hub
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`solidity dimension (to a value greater than 1) to maintain the “blade taper” and
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`“hub tip” ratios disclosed in Davies. UTC-2015, ¶ 72. Although Davies discloses
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`that particular blade taper ratio and hub tip ratio were chosen for the design (GE-
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`GE v. UTC
`IPR2016-00952
`GE-1034.009
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`1005.007), a person of ordinary skill in the art would recognize that there are a
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`range of values for each of these parameters that would be suitable for a low
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`pressure ratio variable pitch fan (e.g., Davies). A 1971 report published by Rolls
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`Royce, for example, discloses that a low pressure ratio variable pitch fan can
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`utilize a hub to tip ratio ranging from 0.5 to 0.6.
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`GE-1030.006, Fig. 9 (annotations in red)
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`13. For a hub tip ratio of 0.55 and blade tip solidity of 0.89,3 a person of
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`ordinary skill would recognize that the blade taper ratio as defined in Davies4 could
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`3 The blade tip solidity value of 0.89 is based on modifying the tip chord dimension
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`of the fan blade disclosed in Davies from 10 inches to 12 inches. See GE-1003, ¶
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`95.
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`GE v. UTC
`IPR2016-00952
`GE-1034.010
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`solidity of 0.95 based on the equation set forth in Davies:
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`be maintained at 1.7 (the value disclosed in Davies) with a corresponding hub
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`(cid:1846)(cid:1861)(cid:1868) (cid:1845)(cid:1867)(cid:1864)(cid:1861)(cid:1856)(cid:1861)(cid:1872)(cid:1877)
`(cid:1846)(cid:1853)(cid:1868)(cid:1857)(cid:1870) (cid:1844)(cid:1853)(cid:1872)(cid:1861)(cid:1867)(cid:3404)
`(cid:1834)(cid:1873)(cid:1854) (cid:1845)(cid:1867)(cid:1864)(cid:1861)(cid:1856)(cid:1861)(cid:1872)(cid:1877)∗(cid:1834)(cid:1873)(cid:1854) (cid:1846)(cid:1861)(cid:1868) (cid:1844)(cid:1853)(cid:1872)(cid:1861)(cid:1867)
`0.89
`(cid:1846)(cid:1861)(cid:1868) (cid:1845)(cid:1867)(cid:1864)(cid:1861)(cid:1856)(cid:1861)(cid:1872)(cid:1877)
`(cid:1846)(cid:1853)(cid:1868)(cid:1857)(cid:1870) (cid:1844)(cid:1853)(cid:1872)(cid:1861)(cid:1867)∗(cid:1834)(cid:1873)(cid:1854) (cid:1846)(cid:1861)(cid:1868) (cid:1844)(cid:1853)(cid:1872)(cid:1861)(cid:1867)(cid:3404)
`1.7∗0.55 (cid:3404)0.95
`(cid:1834)(cid:1873)(cid:1854) (cid:1845)(cid:1867)(cid:1864)(cid:1861)(cid:1856)(cid:1861)(cid:1872)(cid:1877)(cid:3404)
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`A hub solidity of 0.95 (less than 1) would enable the fan blade disclosed in Davies
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`to reverse pitch using either the fine pitch or coarse pitch mode. GE-1005.006.
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`14.
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`I also disagree with Dr. Mathioudakis that the chord dimension could
`
`not be increased due to the required gap between the fan rotor and stator vanes—
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`two rotor chord lengths. UTC-2015, ¶ 75. A person of ordinary skill would be
`
`aware of available design modifications, such as adjusting the length of the bypass
`
`duct to re-position the core inlet so that it is not obstructed by the stators. This
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`would enable an increase in chord tip dimension, while allowing for the same
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`rotor/stator spacing without obstructing the core inlet. In addition, due to the high
`
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`4 Blade taper ratio more commonly refers to the ratio of the blade cross-sectional
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`area distribution from tip to hub, which is a function of the chord dimension and
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`the wall thickness of the blade. In contrast, Davies defines taper ratio as a function
`
`of just the chord dimension (tip to hub).
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`GE v. UTC
`IPR2016-00952
`GE-1034.011
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`stagger angle near the blade tip, changes in the chord dimension of the fan blade
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`near the tip would result in only a small increase in the axial length that would
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`need to be accommodated.
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`V. DEFINITION OF “BYPASS FLOW PASSAGE PRESSURE RATIO”
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`15. Claim 1 of the 412 Patent requires a bypass flow passage having a
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`pressure ratio that is between 1.1 and 1.35 with regard to an inlet pressure and an
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`outlet pressure of said bypass flow passage. According to Dr. Mathioudakis, the
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`inlet pressure of the bypass flow passage is to be measured at the forward most
`
`portion of the fan nacelle (i.e., the inlet to the engine). See UTC-2015, ¶ 93. For
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`the following reasons, I disagree.
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`16. The 412 Patent consistently describes that the inlet to the bypass flow
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`passage is co-extensive with the fan. Claim 1, for example, describes that the
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`propulsor (i.e., fan) is located at the inlet of the bypass flow passage. See GE-1001
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`at claim 1 (“the propulsor is located at an inlet of a bypass flow passage”); see also
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`GE-1001 at 1:44-45, 2:49-50. Furthermore, the figures of the 412 Patent illustrate
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`that the bypass flow passage begins at the fan. As shown below, Figure 1 indicates
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`that the bypass flow passage B and core flow passage C are located downstream of
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`the fan. The area located upstream of the fan is not marked as part of the bypass
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`GE-1034.012
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`flow passage (B). That is because the flow path upstream of the fan is the inlet
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`flow path, and not part of the bypass flow passage.
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`GE-1001.002, Figure 1 (annotations in red)
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`17. A person of ordinary skill would also understand that Figure 1 of the
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`412 Patent does not illustrate the forward portion of the fan nacelle. Figure 1
`
`illustrates the fan case because it immediately surrounds the fan blades, and is not
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`aerodynamically shaped for receiving inlet flow, as the front of the fan nacelle
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`would be. A prior art fan case and fan nacelle (US Patent Application Publication
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`No. 2009/0314881) are reproduced below to illustrate this distinction. As shown
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`below, a fan case is co-extensive with the fan blades, while the front of the fan
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`GE v. UTC
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`GE-1034.013
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`nacelle is forward of the fan blades and is aerodynamically shaped to receive inlet
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`flow:
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`GE-1024.002, Figure 1A (annotations in red)
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`VI. BYPASS FLOW PASSAGE PRESSURE RATIO IN DAVIES
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`18. Dr. Mathioudakis contends that the structures located in the bypass
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`flow passage of the M45SD-02 engine disclosed in Davies could cause pressure
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`losses that exceed 2%-7%. UTC-2015, ¶ 94. I disagree with Dr. Mathioudakis’
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`analysis, as described below.
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`19. As I explained in my initial declaration, the fan pressure ratio value
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`disclosed in Davies differs from a bypass flow passage pressure ratio as claimed in
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`the 412 Patent based on the location of the outlet pressure used in the calculation.
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`GE v. UTC
`IPR2016-00952
`GE-1034.014
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`GE-1003, ¶ 76. The fan pressure ratio disclosed in Davies utilizes a total pressure
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`measured at location P4 as the numerator, while claim 1 of the 412 Patent uses a
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`total pressure measured at P2 as the numerator for the bypass flow passage
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`pressure ratio. Accordingly, the fan pressure ratio differs from the claimed bypass
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`flow passage pressure ratio based on the total pressure losses that occur between
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`P2 and P4 (i.e., the bypass flow passage pressure ratio will be less than the fan
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`pressure ratio). This means that structures that are not located between P2 and P4
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`shown below are irrelevant to assessing the difference between a fan pressure ratio
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`and bypass flow passage pressure ratio because those structures would not cause
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`pressure losses between P2 and P4.
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`GE-1003, ¶ 76
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`GE v. UTC
`IPR2016-00952
`GE-1034.015
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`A.
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`Features Not Located Between P2 and P4 Are Irrelevant to
`Assessing the Difference Between the Fan Pressure Ratio and
`Bypass Flow Passage Pressure Ratio for a Given Engine
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`20. First, Dr. Mathioudakis identifies several features associated with the
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`engine inlet, including an “elongated and divergent inlet cowl, an annular inlet
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`constriction on the inlet cowl, an open-nose fan hub, [and] acoustic treatments….”
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`UTC-2015, ¶ 94. These features are all located upstream of the bypass flow
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`passage (not between P4 and P2), and therefore would not create pressure losses
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`that would contribute to the difference between a fan pressure ratio value and
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`corresponding bypass flow passage pressure ratio:
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`UTC-2015, ¶ 90 (annotations by Dr. Abhari in red)
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`GE-1034.016
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`21. Second, Dr. Mathioudakis identifies “an elongated core cowl.” UTC-
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`2015, ¶¶ 86, 94. The elongated core cowl is clearly downstream of the bypass flow
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`passage outlet (to the right of P2), and therefore would not cause a pressure loss in
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`the bypass flow passage:
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`UTC-2015, ¶ 86 (annotations by Dr. Abhari in red)
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`22. Third, Dr. Mathioudakis identifies several features associated with the
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`stator vanes, including their shape, quantity, and a constriction in the duct near the
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`stator vanes. UTC-2015, ¶¶ 82-83, 87. Because these features are located between
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`P3 and P4 (not between P4 and P2), the pressure losses caused by these structures
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`GE-1034.017
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`are encompassed in the fan pressure ratio value disclosed in Davies, and thus
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`would not contribute to a difference between the fan pressure ratio and the
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`corresponding bypass flow passage pressure ratio for the M45SD-02.
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`GE-1005.019, Figure 1 (annotations in red)
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`23. The figure below summarizes the structures identified by Dr.
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`Mathioudakis, which are not located between P4 and P2, and therefore do not
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`contribute to any difference between the fan pressure ratio and bypass flow
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`passage pressure ratio:
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`GE-1034.018
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`GE-1005.019, Figure 1 (annotations in red)
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`B.
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`Features of a Conventional Bypass Duct Disclosed in Davies
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`24. Dr. Mathioudakis also identifies several structures as
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`
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`“unconventional,” including a compressor bleed valve, acoustic linings, an
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`elongated bypass flow passage, and constricting outlet nozzle. UTC-2015, ¶¶ 82-
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`83. I disagree with Dr. Mathioudakis that these features of the M45SD-02 are
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`unconventional.
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`25. First, bleed ports/valves are standard components of a turbofan
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`engine. U.S. Patent No. 3,898,799, which issued in 1975, describes a compressor
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`bleed valve that can discharge either into the bypass duct or be routed through a
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`GE-1034.019
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`strut to the atmosphere. GE-1025 at 1:13-20, 2:51-55 (“a portion of the air
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`supplied by intermediate pressure compressor 2 can be bled off by the apparatus of
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`the present invention either into bypass duct 12…or via hollow struts 16, passing
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`through bypass duct 12, to atmosphere….”). The presence of bleed valves in
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`turbofan engines is also acknowledged in an exhibit submitted by Patent Owner in
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`this proceeding. UTC-2007.006 at Figure 10 (showing thrust loss of
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`approximately 2% for “bleeds/leakage”).
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`26. Second, acoustic linings on the walls of the bypass duct are also a
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`feature of a typical turbofan engine. GE-1026.007 (“Currently acoustic liners are
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`used in the inlet, fan case, aft bypass duct, and core nozzle to attenuate both fan
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`and core engine noise.”); see also UTC-2007.006 (“Shortening the inlet and
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`exhaust also reduces the area available for the acoustic liner used for noise
`
`attenuation.”).
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`27. Third, Dr. Mathioudakis contends that the length of the bypass flow
`
`passage could cause significant pressure losses. UTC-2015, ¶ 82, 88, 94. The text
`
`of Davies does not describe the bypass flow passage as elongated. See GE-1005.
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`A person of ordinary skill in the art would understand that the length of the bypass
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`duct for a turbofan engine can vary based on the desired engine characteristics,
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`such as noise, performance, and installation concerns. As shown below, Davies
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`GE-1034.020
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`illustrates that the outlet to the bypass flow passage is co-extensive with the turbine
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`exit case of the low pressure turbine:
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`GE-1005.019, Figure 1 (annotations in red)
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`The length of the bypass flow passage in Davies is only slightly longer than those
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`of some other commercial turbofan engines. For example, the bypass flow passage
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`outlet of the GE-90 turbofan engine, which has been in commercial service since
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`the 1990s, is co-extensive with the middle of the low pressure turbine. The
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`M45SD-02 bypass flow passage is therefore only slightly longer:
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`GE-1014.028, Figure 1-8e (annotations in red)
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`28. Fourth, I disagree with Dr. Mathioudakis that a constricting outlet
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`nozzle is unconventional. UTC-2015, ¶ 94. Davies states that the “fan duct exit
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`nozzle has approximately 70% of the fan flow area.” UTC-2015, ¶ 86 (citing GE-
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`1005.009). A person of ordinary skill in the art would understand that the fan flow
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`area must be sized to provide airflow for both the bypass flow passage and the
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`engine core, while the bypass duct nozzle is only sized to discharge the airflow
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`from the bypass flow passage. Accordingly, the bypass duct nozzle exit area will
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`necessarily be less than the fan flow area. Furthermore, nozzles used in the bypass
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`ducts of turbofan engines for commercial aircraft are convergent (i.e., have a
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`constricting area), which is required to increase the velocity of the flow to generate
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`thrust. GE-1014.060.
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`29.
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`I also disagree with Dr. Mathioudakis that the constricting nozzle
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`would result in “substantial changes in pressure” that are relevant to assessing the
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`difference between fan pressure ratio and bypass flow passage pressure ratio.
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`UTC-2015, ¶¶ 25, 86. As I explained during my deposition, there is a difference
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`between a change in static pressure and a change in total pressure with regards to
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`air flowing through a nozzle. UTC-2013 at 89:6-13. While a nozzle may cause a
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`large change in static pressure, a person of ordinary skill would recognize that the
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`change in total pressure would be small (~1-2%). UTC-2013 at 89:24-90:6. Total
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`pressure (i.e., stagnation pressure) is the relevant parameter for calculating fan
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`pressure ratio, and assessing the difference between fan pressure ratio and bypass
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`flow passage pressure ratio. GE-1001 at 2:55-58.
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`
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`C. Additional Features in the Davies Bypass Flow Passage
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`30.
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`In my opinion, there are only two features of the M45SD-02 located
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`between P4 and P2 in the bypass flow passage, which are not conventional
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`components of a bypass duct—the auxiliary flow intakes and core inlet located
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`behind the stator vanes. In my opinion, these features would typically cause
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`minimal pressure losses.
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`31.
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`I disagree with Dr. Mathioudakis that the auxiliary flow intakes used
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`for reverse thrust would always create gaps and have associated mechanical
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`structures that are exposed to the bypass flow passage. UTC-2015, ¶ 82. A person
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`of ordinary skill in the art would understand that auxiliary flow intakes used for
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`reverse thrust are aerodynamically designed to provide a continuous inner surface
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`in the bypass duct when closed, and also would not include mechanical structures
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`exposed to the bypass flow passage, as shown below. See e.g., GE-1028 at 2:38-
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`52.
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`GE-1028.003, Figure 2
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`
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`32.
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`I also disagree with Dr. Mathioudakis that the core splitter nose in
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`Davies would create a significant pressure loss that would result in the total
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`pressure losses in the bypass flow passage to exceed 7%. UTC-2015, ¶¶ 82, 94.
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`Although it is not typical for the core inlet to be located aft of the fan stators in a
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`turbofan engine, it is my opinion that the total pressure losses associated with this
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`configuration would be minimal. Davies describes the core splitter nose as
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`“relatively blunt.” GE-1005.009. A typical core splitter nose is aerodynamically
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`designed only for forward airflow, and therefore has a sharp contour. GE-
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`1029.004 (“The Q-fan T-55 splitter lip is more rounded than the sharp lip of the
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`QCSEE model which would suggest higher losses for the QCSEE model.”). In
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`contrast, a core splitter nose for an engine that utilizes a variable pitch fan to
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`produce reverse thrust (e.g., Davies) must be aerodynamically shaped for both
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`forward airflow and reverse airflow. See e.g., GE-1005.009 (“This, together with
`
`the relatively blunter splitter nose, enables the reverse flow to enter the core with
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`minimum loss.”). Accordingly, the core splitter nose of the M45SD-02 would be
`
`thicker and have a slightly larger curvature shape than a typical sharp splitter nose.
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`GE-1029.004. Nonetheless, one of ordinary skill would understand that it would
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`be aerodynamically shaped to minimize losses during forward thrust operation.
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`D. UTC-2007 Confirms My Opinion That Losses in The Bypass Flow
`Passage Would Not Exceed 7%
`
`33. Dr. Mathioudakis has cited a figure in exhibit UTC-2007 to allegedly
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`
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`show that the losses in the bypass flow passage can exceed 7%. UTC-2015, ¶ 43.
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`UTC-2007 discloses the plot below of net thrust loss v. fan pressure ratio for a
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`turbofan engine. A proper reading of this figure and under the proper
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`interpretation of bypass flow passage pressure ratio, this plot actually confirms my
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`opinion that the losses in the bypass flow passage will be less than 7%.
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`UTC-2007.006, Figure 10 (annotations in red)
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`
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`During his deposition, Dr. Mathioudakis stated that percentage net thrust loss
`
`corresponds to pressure loss approximately 1:1. GE-1031 at 39:14-22. I have
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`adopted the 1:1 correspondence described by Dr. Mathioudakis.
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`34.
`
`I disagree with Dr. Mathioudakis that exhibit UTC-2007 demonstrates
`
`that the relevant bypass flow passage pressure losses would exceed 7%. Under the
`
`correct definition of bypass flow passage pressure ratio, the losses associated with
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`the inlet, fan rotor (i.e., fan disk and blades), and fan static structure (i.e., exit guide
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`vanes) shown in UTC-2007 would not contribute to any difference between the fan
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`pressure ratio and corresponding bypass flow passage pressure ratio because those
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`structures are not located between P4 and P2, as shown below:
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`GE-1005.019, Figure 1 (annotations in red)
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`
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`The difference between fan pressure ratio and a corresponding bypass flow passage
`
`pressure ratio would be due to the pressure losses associated with the duct/nozzle
`
`and bleeds/leakage. As shown in Figure 10 of UTC-2007 (see above), the pressure
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`loss due to the duct/nozzle is approximately 5%, while the pressure loss due to the
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`bleeds/leakage is less than 2%. This results in a total pressure loss in the bypass
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`flow passage of less than 7%, which is consistent with the analysis I provided in
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`my initial declaration. See GE-1003, ¶ 77. Thus, a person of ordinary skill in the
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`art would understand that based on a fan outer pressure ratio of 1.27, as disclosed
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`in Davies (GE-1005.005), the corresponding bypass flow passage pressure ratio
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`would be no less than 1.18 (i.e., 1.27 x 0.93 = 1.18).
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`35. Even under Patent Owner’s proposed construction of bypass flow
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`passage pressure ratio (i.e., using the forward most point of the fan nacelle as the
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`inlet), the corresponding bypass flow passage pressure ratio for a fan outer pressure
`
`ratio of 1.27 would fall within the claimed range (1.1 to 1.35). In particular, Figure
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`10 of UTC-2007 discloses that the total pressure losses for all of the structures
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`would be approximately 9%,5 which yields a bypass flow passage pressure ratio of
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`1.13 (i.e., 1.27 * 0.91 = 1.16).
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`5 As I described above, the 2% pressure losses from the stators (i.e., static
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`structure) shown in UTC-2007 are already encompassed in the fan pressure ratio
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`value, and therefore do not contribute to the difference between fan pressure ratio
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`and bypass flow passage pressure ratio.
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`VII.
`
`JURAT
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`
`
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`36.
`
`I declare that all statements made herein of my own knowledge are
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`true and that all statements made on information and belief are believed to be true;
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`and further that these statements were made with the knowledge that willful false
`
`
`statements and the like so made are punishable by fine or imprisonment, or both,
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`under Section 1001 of Title 18 of the United States Code.
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`I declare under penalty ofperjury that the foregoing is true and
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`37.
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`
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`correct.
`Executed on i ?'l’ "° ', 2017 a g . XLAK/K—OA
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`
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`Reza Abhari, PhD.
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`I
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`|PR2016-00952
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