`
`BEFORE THE PATENT TRIAL AND APPEAL BOARD
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`
`
`In re U.S. Patent No. 8,678,743
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`
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`Filed:
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`February 4, 2013
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`Issued:
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`March 25, 2014
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`Inventors: William G. Sheridan, Karl L. Hasel
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`Assignee: United Technologies Corporation
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`Title:
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`Method for Setting a Gear Ratio of a Fan Drive Gear System of a Gas
`Turbine Engine
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`Mail Stop PATENT BOARD, PTAB
`Patent Trial and Appeal Board
`U.S.P.T.O.
`P.O. Box 1450
`Alexandria, VA 22313-1450
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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. 8,678,743 (“the 743
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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-1003.001
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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. 8,678,743
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`(the “743 Patent”) to William G. Sheridan et al.
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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:
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`GE-1001 U.S. Patent No. 8,678,743 to William G. Sheridan et al.
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`GE-1002 Prosecution File History of U.S. Patent No. 8,678,743.
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`GE-1005 Declaration of Raymond Drago
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`GE-1006 Curriculum Vitae of Raymond Drago
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`GE-1007 Bruce E. Wendus et al., Follow-On Technology Requirement Study
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`for Advanced Subsonic Transport (August 2003).
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`GE-1009 Cesare A. Hall et al., Engine Design Studies for a Silent Aircraft,
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`Journal of Turbomachinery (2007).
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`GE-1010 U.S. Patent No. 7,021,042 B2 to Law, Geartrain Coupling for a
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`Turbofan Engine (issued April 4, 2006) (“Law”).
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`GE-1011 William S. Willis, Quiet Clean Short-Haul Experimental Engine
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`(QCSEE) Final Report (August 1979).
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`2
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`GE-1003.002
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`
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`GE-1012 Bill Sweetman et al., Pratt & Whitney’s surprise leap, INTERAVIA
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`(June 1998).
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`GE-1013 Gerald Brines, The Turbofan of Tomorrow, Mechanical Engineering
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`(August 1990) (“Brines”).
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`GE-1014 Excerpts from Jack D. Mattingly, Elements of Gas Turbine
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`Propulsion (1996).
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`GE-1015 Bill Gunston, Pratt & Whitney PW8000, Jane’s Aero-Engines Issue 7
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`(March 2000).
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`GE-1017 Richard Whitaker, ALF502: plugging the turbofan gap, Flight
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`International (Jan. 30, 1982).
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`GE-1019 Excerpts from Prosecution File History of U.S. Patent Application
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`No. 14/705,459.
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`GE-1020 Excerpts from Prosecution File History of U.S. Patent Application
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`No. 14/179,827.
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`GE-1022
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`Joachim Kurzke, Preliminary Design, Aero-Engine Design: From
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`State of the Art Turbofans Towards Innovative Architectures (March
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`3-7, 2008).
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`3
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`GE-1003.003
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`3.
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`In forming my opinions expressed below, I have considered the
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`documents listed above; and my knowledge and experience based upon my work
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`in this area as described below.
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`4.
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`The application that led to the issuance of the 743 Patent was filed on
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`February 4, 2013. I am familiar with the technology at issue and am aware of the
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`state of the art around this time. Based on the technology disclosed in the 743
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`Patent, a person of ordinary skill in the art (“POSITA”) would include someone
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`who has a M.S. degree in in Mechanical Engineering or Aerospace Engineering as
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`well as at least 3-5 years of experience in the field of gas turbine engine design and
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`analysis. My analyses and opinions below are given from the perspective of a
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`POSITA in these technologies in this timeframe, unless stated otherwise.
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`II. QUALIFICATIONS AND COMPENSATION
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`5.
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`I am currently a Full Professor of Aerothermodynamics at the Swiss
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`Federal Institute of Technology (“ETH”) in Zurich, Switzerland, which is a
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`position I have held since 1999. I am also the head of the Laboratory for Energy
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`Conversion at ETH.
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`6.
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`I received a BA degree in Engineering Science from Oxford
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`University in 1984, and a PhD from the Aeronautical and Astronautical
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`Engineering Department at the Massachusetts Institute of Technology (“MIT”) in
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`1991.
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`4
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`GE-1003.004
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`7. My research in the field of gas turbine technology began in 1984 at
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`Oxford University and continued throughout my academic career at Oxford and at
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`MIT. I began working with the relevant technology in the commercial industry in
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`1991. From 1991-1994, I was a Senior Research and Development Engineer for
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`Textron Lycoming in Stratford CT, where I focused on research, development and
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`design of engine components for next generation commercial and military gas
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`turbine engines for aircraft propulsion. From 1994-1995, I was the Section Head
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`of Propulsion and Energy Research at the Calspan Advanced Technology Center in
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`Buffalo NY, where I was responsible for heading the group performing research
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`and development of gas turbine technology related to turbine and combustor
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`performance and reliability as well as overall engine operability in severe
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`environments.
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`8.
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`In 1995, I became an Assistant Professor in the Aeronautical
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`Engineering Department at the Ohio State University in Columbus Ohio, with a
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`joint appointment in the Mechanical Engineering Department. In 1997, I was
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`promoted to Associate Professor with Tenure at the Ohio State University, where I
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`remained until 1999 when I received my current position at ETH. From 1995-
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`1999, I was also one of the two founders and the Associate Director of the Gas
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`Turbine Laboratory at the Ohio State University.
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`5
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`GE-1003.005
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`9.
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`In 1999, I became the Full Professor and Director of the
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`Turbomachinery Laboratory at ETH. The Turbomachinery Laboratory at ETH was
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`founded in 1892 and is one of the oldest university research centers performing
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`long-term research in turbomachinery, including gas turbine technology. In 2008,
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`the name of the laboratory was changed from Turbomachinery Lab to Laboratory
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`for Energy Conversion (LEC) to better reflect the breadth of the research.
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`10.
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`I am a member of the Swiss Academy of Engineering Sciences, and I
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`am also a fellow of the American Society of Mechanical Engineers (ASME). I
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`have been a member of the ASME International Gas Turbine Institute (IGTI)
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`Turbomachinery Committee since 1995. From 2004-2010, I was a member of the
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`Board of Directors of the ASME IGTI, as well as its Chairman from 2008 to 2009.
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`11.
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`I have the honor of being the recipient of the 2014 R. Tom Sawyer
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`Award of ASME for “significant contributions to the gas turbine industry in both
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`the U.S. and Europe, and for exemplary service to the IGTI”. This is one of the
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`highest awards in the field of gas turbine technology.
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`12.
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`I have written about and studied the field of gas turbine engine design
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`extensively for about three decades. I have taught aircraft engine design for about
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`20 years, supervised over 200 MSc and PhD theses and have been the author of
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`over 250 technical papers, many of which relate to gas turbine engine component
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`6
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`GE-1003.006
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`performance, reliability and design. A listing of my technical papers is included in
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`my curriculum vitae, which is attached as GE-1004.
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`13.
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`I am being compensated at an hourly rate of 500.00 Swiss francs
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`(CHF) for work performed in Switzerland, and 583.00 CHF for work performed in
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`the United States. These are my standard hourly rate for consulting engagements.
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`My compensation is not dependent on the substance of my statements in this
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`Declaration.
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`I.
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`RELEVANT LEGAL STANDARDS
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`14.
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`I have been asked to provide my opinions regarding whether the
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`claims of the 743 Patent are anticipated or rendered obvious by the prior art.
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`15.
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`I have been informed that in order for prior art to anticipate a claim
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`under 35 U.S.C. § 102, the reference must disclose every limitation of the claim.
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`16.
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`I have been informed that a claimed invention is not patentable under
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`35 U.S.C. § 103 if the differences between the invention and the prior art are such
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`that the subject matter as a whole would have been obvious at the time the
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`invention was made to a POSITA. I also understand that the obviousness analysis
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`takes into account factual inquiries including the level of ordinary skill in the art,
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`the scope and content of the prior art, the differences between the prior art and the
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`claimed subject matter, and any secondary considerations which may suggest the
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`claimed invention was not obvious. I have been informed that a claim can be
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`7
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`GE-1003.007
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`obvious in light of a single prior art reference or multiple prior art references. I
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`understand that a claim can be obvious in light of a single reference if a motivation
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`exists, such as common sense or knowledge of one of skill in the art, to supply any
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`missing limitations of the claim to that reference.
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`17.
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`I have been informed by legal counsel that the Supreme Court has
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`recognized several rationales for combining references or modifying a reference to
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`show obviousness of claimed subject matter. I understand some of these rationales
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`include the following: combining prior art elements according to known methods
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`to yield predictable results; simple substitution of one known element for another
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`to obtain predictable results; use of a known technique to improve a similar device
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`(method, or product) in the same way; applying a known technique to a known
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`device (method, or product) ready for improvement to yield predictable results;
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`choosing from a finite number of identified, predictable solutions, with a
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`reasonable expectation of success; and some teaching, suggestion, or motivation in
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`the prior art that would have led a POSITA to modify the prior art reference or to
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`combine prior art reference teachings to arrive at the claimed invention.
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`II. BACKGROUND OF THE TECHNOLOGY
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`18. The following paragraphs regarding turbofan architecture and design
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`are based on prior art to the 743 Patent.
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`8
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`GE-1003.008
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`A. Turbofan Engines
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`19. Turbofan engines are a type of gas turbine engine commonly used in
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`commercial aviation. These engines are often represented in publications by a
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`variety of images depending upon the level of detail required, including
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`photographs, cut-away images, and cross-sections. Below is an illustration of the
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`GE90 high-bypass ratio turbofan engine used in subsonic aircraft. The GE90 was
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`introduced in 1995 and powers the Boeing 777.
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`GE-1014.028, Figure 1-8e.
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`20. The illustration below is another representation of the GE90 without
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`the fan nacelle and a portion of the fan case cut-away.
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`9
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`GE-1003.009
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`GE90 Turbofan Cutaway Image
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`21. Turbofan engines are generally comprised of the following sections:
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`an inlet section, a fan section, a compressor section, a combustor section, a turbine
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`section, and an exhaust section. GE-1014.024-.027, Figure 1-7. The compressor
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`section typically includes a low pressure compressor (i.e., LP Compressor, LPC, or
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`booster) and a high pressure compressor (i.e., HP Compressor or HPC). GE-
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`1014.024, Figure 1-7. Similarly, the turbine section typically includes a low
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`pressure turbine (i.e., LP Turbine or LPT) and a high pressure turbine (i.e., HP
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`Turbine or HPT). GE-1014.024, Figure 1-7. These sections of a conventional
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`turbofan engine are shown in the cross-section figure below from a 1996 textbook
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`(as well as in the images above of the GE90 engine). The schematic cross section
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`10
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`GE-1003.010
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`is very similar to the cutaway image but illustrates the major structural components
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`and architecture schematically. A person of ordinary skill in the art would be very
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`familiar with schematic cross section drawings of turbofan engines and would be
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`able to understand major structural components and architecture of an engine from
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`a schematic cross section.
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`GE-1014.024, Figure 1-7 (annotations in color)
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`22. As shown above, air enters a turbofan engine through the inlet and
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`then the fan section. GE-1014.023-.024; GE-1013.006-.007. After passing
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`through the fan section, the air travels via one of two flow paths: (1) the bypass
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`flow path (colored blue) or (2) the core flow path (colored orange). GE-1014.024,
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`11
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`GE-1003.011
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`.034 (“The thrust for a turbofan engine with separate exhaust streams is equal to
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`the sum of the thrust from the engine core FC and the thrust from the bypass stream
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`FB.”). It is standard practice to refer to the ratio of the mass flow rate of air
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`bypassing the engine core (i.e., the bypass flow) to the mass flow rate of air
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`passing through the engine core (i.e., the core flow) as the bypass ratio. GE-
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`1014.034 (“The bypass ratio …is the ratio of the mass flow through the bypass
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`stream to the core mass flow….”). In current commercial aircraft engines the
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`bypass ratio can range anywhere from approximately 5 to 8.5, meaning that most
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`of the air that enters the engine travels through the bypass flow path. GE-
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`1014.145.
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`23. Because higher bypass ratio correlates to improved propulsive
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`efficiency, it is generally preferable for a turbofan engine to have a higher bypass
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`ratio. GE-1013.005 (“The higher the ratio of bypassed air to air passing through
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`the engine, the greater the fuel efficiency of the engine. The need for such engines
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`has been spurred by increasing airplane traffic, which raises noise, environmental,
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`and fuel consumption issues.”).
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`24. The bypass flow travels through the bypass duct and then exits the
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`engine from the exhaust section to generate thrust. GE-1014.024, .034. The core
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`flow, on the other hand, travels through the compressor section, combustor section,
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`and turbine section before exiting a turbofan engine via the exhaust section. GE-
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`12
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`GE-1003.012
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`1014.023-.024, .053-.060. The core flow is first compressed by the low pressure
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`compressor and then the high pressure compressor. GE-1014.023-.024, .055 (“The
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`function of the compressor is to increase the pressure of the incoming air.”). As
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`represented in the figure below, both the LPC and HPC include multiple stages1
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`with rotating blades. Each stage compresses the air, forcing the pressure and
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`temperature of the air to rise.
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`GE-1014.024, Figure 1-7(annotations in color)
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`1 A stage consists of a rotating disk (i.e., rotor disk) that holds a plurality of blades,
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`and a set of stationary airfoils known as stator vanes. As described below, each
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`stage can be identified by a vertical line that protrudes from the centerline axis,
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`which represents the rotor disk of a stage.
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`13
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`GE-1003.013
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`25. After exiting the high pressure compressor, the core flow enters the
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`combustor section, where it is mixed with fuel and ignited, increasing the
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`temperature of the gas mixture and raising the energy level of the gas. GE-
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`1014.023-.024, .057 (“The combustor is designed to burn a mixture of fuel and air
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`and to deliver the resulting gases to the turbine”). The combustor section is
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`highlighted in the figure below.
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`GE-1014 at 10, Figure 1-7 (annotations in color)
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`26. Once the core flow is compressed and heated, it must be made to do
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`
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`useful work. After exiting the combustor section, the core flow is then expanded
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`through the high pressure turbine, which drives the high pressure compressor via
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`14
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`GE-1003.014
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`the high spool shaft. GE-1014.023-.024, .057, .088 (“The high-pressure turbine
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`drives the high-pressure compressor….”). “The assembly containing the high-
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`pressure turbine, high-pressure compressor, and connecting shaft is called the high-
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`pressure spool,” which is also referred to as the high speed spool or high spool.
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`GE-1014.088 (emphasis in original). The high pressure turbine and high spool
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`shaft are highlighted in the figure below.
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`GE-1014.024, Figure 1-7 (annotations in color)
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`27. The core flow is then further expanded across the low pressure
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`turbine, which drives the low pressure compressor and the fan via the low spool
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`shaft. GE-1014.023-.024, .057, .088 (“…the low-pressure turbine drives the
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`15
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`GE-1003.015
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`fan…and low-pressure compressor….”). “[The assembly] containing the low-
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`pressure turbine, fan or low-pressure compressor, and connecting shaft is called the
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`low-pressure spool,” which is also referred to as the low speed spool or low spool.
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`GE-1014.088 (emphasis in original). The low pressure turbine and low spool shaft
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`are highlighted in the figure below.
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`GE-1014.024, Figure 1-7 (annotations in color)
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`28. The above engine configuration is commonly called a two-spool
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`turbofan engine. GE-1009.010; GE-1014.106-.107. The major commercial
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`16
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`GE-1003.016
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`aircraft propulsion gas turbine manufacturers generally use a conventional two
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`spool configuration2 with co-axial high and low spools as shown below:
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`GE-1014.024, Figure 1-7 (annotations in red)
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`The low spool shaft is depicted above in blue along with the other low spool
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`components (i.e.¸ the stages of the low pressure compressor and low pressure
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`turbine), while the high spool shaft is depicted above in red along with the other
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`high spool components (i.e., the stages of the high pressure compressor and high
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`pressure turbine). Each stage of a particular section of the engine (e.g., the low
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`2 Rolls Royce typically utilizes a turbofan configuration that includes three spools.
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`17
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`GE-1003.017
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`pressure compressor) can be identified by a vertical line that protrudes from the
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`centerline axis, which represents the rotor disc for a stage that is connected to
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`either the low spool shaft or high spool shaft. For example, a person of ordinary
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`skill in the art would recognize that the high pressure compressor illustrated above
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`includes 5 stages, while the high pressure turbine includes 1 stage.
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`B.
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`Turbofan Engine Configurations
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`29. Most two-spool turbofan engines in operation today are configured as
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`“direct drive” turbofan engines. In a direct drive turbofan engine, the fan is
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`directly connected to the low spool shaft, such that the low pressure turbine, low
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`pressure compressor, and fan all rotate at the same rotational speed. GE-1014.088.
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`Although a direct drive configuration is commonly implemented, it is well
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`understood that a direct drive turbofan engine has limitations with respect to
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`increasing the size of the fan, which is generally required to increase the bypass
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`ratio. GE-1013.005 (“Ultrahigh bypass turbofans are engines that use a large fan at
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`the front of the engine….”). If the size of the fan increases, there must be a
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`corresponding reduction in fan rotational speed to keep mechanical stresses at
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`acceptable levels, and to minimize engine noise. GE-1015.009 (“As bypass ratio is
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`increased (to improve fuel economy and reduce noise), the rotational speed of the
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`larger fan must fall”); GE-1013.006 (“A 14:1 bypass ratio ungeared low spool
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`would produce a much larger engine, with an eight-stage low-pressure turbine and
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`18
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`GE-1003.018
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`a 10 percent higher fan speed. This would result in a 2 to 3 percent thrust reduction
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`and an increase of about 2 EPNdB in takeoff noise.”).
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`30. Furthermore, in a direct drive turbofan, a slower fan rotational speed
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`requires reducing the speed of the LPT, which requires more stages, longer length
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`and higher weight in the LPT to drive the low speed spool. GE-1015.009 (“The
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`problem of a direct-drive engine is that it is difficult to match the rotational speeds
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`of the turbine and fan. As bypass ratio is increased…the rotational speed of the
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`larger fan must fall, demanding…a large, heavy, and costly multistage
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`turbine….”). Moreover, as the rotational speed of the fan is decreased, the
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`required torque generally increases, requiring a thicker LP shaft. A larger hole in
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`the high pressure turbine rotor disk is required to accommodate a thicker LP shaft,
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`which also increases the stresses on the high pressure turbine rotor significantly.
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`GE-1007.014 (“Similar studies on the high-pressure turbine (HPT) disk also
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`showed an overstressed condition if the bore diameter was increased to
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`accommodate a larger low shaft.”).
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`31. Accordingly, turbofan engine makers have designed and developed
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`alternatives to the two-spool direct drive turbofan engine. One such alternative is
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`generally referred to as a geared turbofan engine. In a geared turbofan engine, a
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`gear arrangement (i.e., a gearbox, gear train, or gear system) is incorporated into
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`the engine between the fan section and the components of the low speed spool.
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`19
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`GE-1003.019
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`GE-1013.006 (“The fan is driven by a high-speed, transonic, LP turbine through
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`a…gear system.”). The gear arrangement is connected to the fan section on one
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`side, and the low spool shaft on the other side. GE-1013.006 (“The most
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`significant feature of the low spool is the gear-driven, variable pitch fan”). The
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`low pressure turbine thus drives the fan section through the gear arrangement, and
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`enables the fan to rotate at a lower rotational speed than the rest of the low speed
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`spool components (such as the low pressure turbine). GE-1015.010 (“The design
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`shaft speeds are 9,000 rpm on the low spool and 3,200 on the fan….”); GE-
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`1013.006 (“The fan is driven by a…LP turbine through a 3:1 reduction ratio
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`planetary gear system.”). Because the fan and low pressure turbine are not directly
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`coupled to one another, the geared configuration enables the fan to operate at its
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`optimal low rotational speed, while the low pressure turbine operates at its optimal
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`high rotational speed. GE-1013.006 (“The 3:1 speed reduction of the fan drive
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`gear system allows the designer to select the fan tip speed for lowest noise and
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`highest fan efficiency, while at the same time maximizing the low spool shaft
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`speed to allow the use of fewer low pressure compressor and turbine stages.”).
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`32. Although geared turbofan engines are not widely used commercially
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`to date, the concept was derived at least as early as the 1970s, and has been
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`extensively developed and tested since then. For example, General Electric
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`20
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`GE-1003.020
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`designed, built, and tested a geared turbofan engine in the 1970s under a contract
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`from NASA:
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`The NASA/GE QCSEE concept is based on a lightweight, high-
`speed, power turbine driving a slower speed, quiet fan. This
`objective required a compatible, compact, lightweight, high-power-
`capability, main reduction gear. Two reduction gears designed,
`manufactured, and rig-tested by Curtiss-Wright under subcontract to
`General Electric have given trouble-free performance throughout the
`engine-demonstration program.3
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`GE-1011.088. Another example of a geared turbofan is the ALF502 that was
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`developed by Lycoming in the 1970s and first commercially sold in the early
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`1980s. The ALF502 included a gear system for the fan having a gear reduction
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`ratio of 2.3:1. GE-1017.005 (“Most of the development work on the 502 has been
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`connected with the fan. Avco decided on a geared fan arrangement….The three
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`planetary gears give a 2.3:1 speed reduction.”). Updated versions of the ALF502
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`are still used today on the British Aerospace 146 regional aircraft and the Avro
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`RJ85, which are used by airlines in Europe.
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`33. The QCSEE geared turbofan engine is illustrated below in a figure
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`from a 1979 report.
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`3 All emphasis has been added unless otherwise noted.
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`21
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`GE-1003.021
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`GE-1011.033, Figure 8 (annotations in red)
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`34.
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`It has been well understood in the aviation industry for decades that a
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`geared turbofan configuration can offer several benefits relative to a direct drive
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`turbofan engine. As explained above, the gearbox decouples the fan from the low
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`pressure turbine and low pressure compressor, which permits the low pressure
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`compressor and low pressure turbine to rotate at a high rotational speed, requiring
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`fewer stages when compared to a conventional direct drive turbofan engine, which
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`can reduce engine weight and maintenance costs. GE-1015.010 (“The [LP
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`Compressor] is not constrained to rotate at the same speed as the fan”); GE-
`
`1013.006 (“The high-speed low spool permits the elimination of a total of three to
`
`22
`
`GE-1003.022
`
`
`
`
`
`five stages cumulatively in the low-pressure compressor and low-pressure
`
`turbine.”).
`
`35. The slow rotational speed of the fan enabled by a geared configuration
`
`also provides several benefits. Fan blade tip speed, which is proportional to fan
`
`rotational speed, is a significant limitation in turbofan engine design. Specifically,
`
`it is well understood that fan blade tip speeds have to be limited to keep noise at an
`
`acceptable level, and to minimize the efficiency losses associated with shock
`
`waves caused by fan blade tip speeds that exceed the speed of sound. GE-
`
`1012.001 (“The fan’s tip speed cannot go much above 450 m/s – above that point,
`
`noise starts to become unacceptable”); GE-1013.006 (“The 3:1 speed reduction of
`
`the fan drive gear system allows the designer to select the fan tip speed for lowest
`
`noise and highest fan efficiency….”); GE-1013.007 (“The low fan-pressure ratio,
`
`wide chord blade has a much lower fan tip speed than a conventional fan (900
`
`versus 1500 feet per second).”). The geared configuration enables a turbofan
`
`engine to have a low fan tip speed even with a high bypass ratio, thereby
`
`minimizing the negative effects associated with supersonic tip speeds. GE-
`
`1012.001-.002; GE-1013.006.
`
`
`
`C. Low Pressure Turbine
`
`36. As noted above, in a turbofan engine the low pressure turbine drives
`
`the low pressure compressor, and also drives the fan directly (i.e., direct drive) or
`
`23
`
`GE-1003.023
`
`
`
`
`
`via a gearbox (i.e., geared fan). GE-1014.088 (“the low pressure turbine drives the
`
`fan…and low-pressure compressor”); GE-1013.006 (“The fan is driven by a high-
`
`speed, transonic, LP turbine through a 3:1 gear reduction ratio planetary gear
`
`system”). The low pressure turbine in a turbofan engine is typically an axial flow
`
`turbine that consists of multiple stages. GE-1014.122 (“The mass flow of a gas
`
`turbine engine…is generally large enough to require an axial turbine…”); GE-
`
`1013.006 (“geared fan with a…three-stage high-speed low-pressure turbine”).
`
`37. One known advantage of a geared turbofan configuration is that it
`
`enables a more compact low pressure turbine section. In a direct drive turbofan,
`
`the rotational speed of the low pressure turbine section is limited because it is
`
`directly connected to the fan section, which means that each stage of the low
`
`pressure turbine can do less work. GE-1012.001-.002 (“As the designers increase
`
`the fan diameter in pursuit of more thrust, the tip speed limit mean an ever lower
`
`rotational speed, so that each blade and stage in the low-pressure turbine can do
`
`less work”). Accordingly, a large number of stages are required to drive the low
`
`pressure compressor and fan section in a direct drive. GE-1012.002 (“The result is
`
`that high-bypass, high-thrust engines have large and complex low-pressure
`
`compressors and turbines”). In a geared turbofan, the low pressure turbine can
`
`rotate faster because its speed is independent of the fan section. GE-1012.002
`
`(“Gearing solves the problem. The LP turbine and compressor spin faster…The
`
`24
`
`GE-1003.024
`
`
`
`
`
`speed of the fan can be optimized for noise, efficiency and structural
`
`considerations”). The result is that each stage of the low pressure turbine in a
`
`geared turbofan can do more work and the low pressure turbine as a whole can be
`
`made smaller in terms of diameter, airfoil count, and stage count. GE-1012.002
`
`(“The LP turbine and compressor spin faster, which means that they can be made
`
`smaller in diameter, shorter and simpler: the engine has 52% fewer compressor and
`
`turbine blades than a conventional turbofan”); GE-1013.006 (“the fan drive gear
`
`system allows…maximizing the low spool shaft speed to allow the use of fewer
`
`low pressure compressor and turbine stages.”).
`
`III. THE 743 PATENT
`
`
`A. Overview of the Claims
`
`
`
`38. The 743 Patent describes a gas turbine engine that incorporates a
`
`planetary fan drive gear system between the fan section and the low pressure
`
`turbine. GE-1001 at 1:30-33. As shown below, the fan drive gear system is
`
`connected to the fan on one side, and the low pressure turbine on the other side via
`
`the shaft of the low speed spool:
`
`25
`
`GE-1003.025
`
`
`
`
`
`GE-1001 at Figure 1 (annotations in red)
`
`
`
`This geared configuration enables the fan to rotate at a different speed than the low
`
`pressure turbine. GE-1001 at 1:30-33. In the claimed embodiments, the low
`
`pressure turbine includes at least three stages and no more than four stages. GE-
`
`1001 at 4:34-36.
`
`39. The 743 Patent also describes the structure of the planetary fan drive
`
`gear system, which is a simple planetary gear train that was well known to a person
`
`of ordinary skill in the art at the time the application that led to the 743 Patent was
`
`filed. The gear ratio of the planetary gear train is between 2.5 and 5.0. GE-1001 at
`
`5:36-39.
`
`40.
`
`In addition to describing a particular engine architecture and gear
`
`system, the 743 Patent also describes several engine parameters of the claimed gas
`
`turbine engine. In particular, the fan blade tip speed of the fan is less than 1400
`
`26
`
`GE-1003.026
`
`
`
`
`
`feet per second (“fps”), and in certain claims is also greater than 1000 fps. GE-
`
`1001 at claims 1, 11, 20. Furthermore, the bypass ratio of the claimed gas turbine
`
`engine is greater than 11.0 and less than 22.0. GE-1001 at claims 1, 20. Finally,
`
`certain dependent claims require a fan pressure ratio of less than 1.7 or less than
`
`1.48. GE-1001 at claims 3, 4.
`
`41. As discussed below, a gas turbine engine with this particular engine
`
`architecture, gear system, and associated engine parameters was well known in the
`
`art before the 743 Patent was filed.
`
`
`
`B. Meaning of Certain Terms in the 743 Patent
`
`42. For a non-expired patent, it is my understanding that a claim subject
`
`to an IPR is interpreted in a manner that is consistent with the broadest reasonable
`
`interpretation in light of the specification. This means that the words of the claim
`
`are given their plain meaning unless that meaning is inconsistent with the
`
`specification. Below I have provided a definition for certain claim terms consistent
`
`with the broadest reasonable interpretation in light of the specification.
`
`1.
`
`“planetary fan drive gear system” and “planetary drive
`gear system”
`
`43. For the purposes of my analysis, I have adopted the construction of
`
`
`
`planetary fan drive gear system articulated by Raymond Drago in his declaration in
`
`support of the petition for inter partes review of the 743 Patent. Accordingly, I
`
`have interpreted a “planetary fan drive gear system” or “planetary drive gear
`
`27
`
`GE-1003.027
`
`
`
`
`
`system” to mean a planetary gear that includes a sun gear, a fixed ring gear
`
`disposed about the sun gear, and a plurality of planet gears supported by a carrier
`
`and positioned between the sun gear and ring gear. See GE-1005 at ¶¶ 33-35.
`
`
`
`2.
`
`“gear ratio”
`
`44. For the purposes of my analysis, I have adopted the construction of
`
`gear ratio articulated by Raymond Drago in his declaration in support of the
`
`petition for inter partes review of the 743 Patent. I have interpreted the phrase
`
`“gear ratio” to mean the ratio of the input rotational speed to the output rotational
`
`speed. See GE-1005 at ¶¶ 36-41.
`
`
`
`3.
`
`“bypass ratio”
`
`45. Claims 1-11 and 20 of the 743 Patent require a bypass ratio between
`
`11.0 and 22.0. It is well known in the art that the term bypass ratio is the ratio of
`
`the mass flow rate of air bypassing the engine core to the mass flow rate of air
`
`passing through the engine core. GE-1014.034 (“The bypass ratio of the
`
`engine…is the ratio of the mass flow through the bypass stream to the core mass
`
`flow….”).
`
`46. The 743 Patent uses the term bypass ratio consistent with its ordinary
`
`meaning in the art. For exam