`
`BEFORE THE PATENT TRIAL AND APPEAL BOARD
`
`
`
`In re U.S. Patent No. 9,121,412
`
`
`
`Filed:
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`July 5, 2011
`
`Issued:
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`September 1, 2015
`
`Inventors: Edward J. Gallagher, Jun Jiang, Becky E. Rose, Jason Elliot, Anthony
`
`
`R. Bifulco
`
`Assignee: United Technologies Corporation
`
`Efficient, Low Pressure Ratio Propulsor for Gas Turbine Engines
`
`
`
`DECLARATION OF REZA ABHARI, PH.D.
`
`Title:
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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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`
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`GE-1003.001
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`
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`
`
`I.
`
`INTRODUCTION
`1.
`
`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
`
`(the “412 Patent”).
`
`2.
`
`In the preparation of this declaration, I have reviewed the relevant
`
`portions of the following documents:
`
`GE-1006
`
`GE-1001 U.S. Patent No. 9,121,412
`
`GE-1002 Prosecution File History of U.S. Patent No. 9,121,412
`
`GE-1005 D.G.M. Davies, et al., A Variable Pitch Fan for an Ultra Quiet
`Demonstrator Engine (1976)
`
`614: VFW’s Jet Feedliner, Flight International (November 4, 1971)
`
`GE-1007 U.S. Patent No. 7,374,403 to Decker et al.
`
`GE-1008 NASA SP-7037 (92), A Cumulative Index to the 1977 Issues of
`Aeronautical Engineering: A Special Bibliography (January 1978)
`(excerpt)
`
`John W. Schaefer et al., Dynamics of High-Bypass-Engine Thrust
`Reversal Using A Variable-Pitch Fan (May 1977).
`
`GE-1010 NASA Technical Reports Server Record Details for GE-1016.
`
`GE-1011 William S. Willis, Quiet Clean Short-Haul Experimental Engine
`(QCSEE) Final Report (August 1979).
`
`GE-1012 Bill Sweetman et al., Pratt & Whitney’s surprise leap, INTERAVIA
`(June 1998).
`
`
`GE-1009
`
`
`
`2
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`GE-1003.002
`
`
`
`
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`GE-1013 Gerald Brines, The Turbofan of Tomorrow, Mechanical Engineering
`(August 1990).
`
`GE-1014 Excerpts from Jack D. Mattingly, Elements of Gas Turbine
`Propulsion (1996).
`
`GE-1015 Bill Gunston, Pratt & Whitney PW8000, Jane’s Aero-Engines Issue 7
`(March 2000).
`
`GE-1016 Bruce E. Wendus et al., Follow-On Technology Requirement Study
`for Advanced Subsonic Transport (August 2003).
`
`GE-1017 Richard Whitaker, ALF502: plugging the turbofan gap, Flight
`International (Jan. 30, 1982).
`
`GE-1018 About NASA Technical Reports Server (www.sti.nasa.gov/find-sti).
`
`GE-1019 University of California at Davis MARC record for Davies
`
`GE-1020 NASA Technical Reports Server Record Details for Schaefer
`
`GE-1021 U.S. 5,141,400 to Murphy et al.
`
`GE-1022 S.A. Savelle et al., Application of Transient and Dynamic
`Simulations to the U.S. Army T55-L-712 Helicopter Engine (1996).
`
`GE-1023 A Summary of Commonly Used Marc 21 Authority Fields, Library of
`Congress
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`
`
`In forming my opinions expressed below, I have considered the
`
`3.
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`documents listed above, and my knowledge and experience based upon my work in
`
`this area as described below.
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`4.
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`The application that led to the issuance of the 412 Patent was filed on
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`July 5, 2011. I am familiar with the technology at issue and am aware of the state
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`
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`3
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`GE-1003.003
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`
`
`
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`of the art around this time. Based on the technology disclosed in the 412 Patent, a
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`person of ordinary skill in the art (“POSITA”) would include someone who has a
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`M.S. degree in in Mechanical Engineering or Aerospace Engineering as well as at
`
`least 3-5 years of experience in the field of gas turbine engine design and analysis.
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`My analyses and opinions below are given from the perspective of a POSITA in
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`these technologies in this timeframe, unless stated otherwise.
`
`II. QUALIFICATIONS AND COMPENSATION
`
`
`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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`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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`
`
`4
`
`GE-1003.004
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`
`
`
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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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`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
`
`for Energy Conversion (LEC) to better reflect the breadth of the research.
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`
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`5
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`GE-1003.005
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`
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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
`
`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.
`
`12.
`
`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
`
`20 years, supervised over 200 MSc and PhD theses and have been the author of
`
`over 250 technical papers, many of which relate to gas turbine engine component
`
`performance, reliability and design. A listing of my technical papers is included in
`
`my curriculum vitae, which is attached as GE-1004.
`
`13.
`
`I am being compensated at an hourly rate of 500.00 Swiss francs
`
`(CHF) for work performed in Switzerland, and 583.00 CHF for work performed in
`
`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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`
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`6
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`GE-1003.006
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`
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`III. 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 412 Patent are anticipated or rendered obvious by the prior art.
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`15.
`
`I have been informed that in order for prior art to anticipate a claim
`
`under 35 U.S.C. § 102, the reference must disclose every element of the claim.
`
`16.
`
`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
`
`that the subject matter as a whole would have been obvious at the time the
`
`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
`
`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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`
`
`7
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`GE-1003.007
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`
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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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`IV. 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 412 Patent.
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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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`
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`8
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`GE-1003.008
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`
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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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`GE90 Turbofan Cutaway Image
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`9
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`
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`GE-1003.009
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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-.026. The compressor section
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`typically includes a low pressure compressor (LP Compressor, LPC, or booster)
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`and a high pressure compressor (HP Compressor or HPC). GE-1014.024, Figure
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`1-7. Similarly, the turbine section typically includes a low pressure turbine (LP
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`Turbine or LPT) and a high pressure turbine (HP Turbine or HPT). GE-1014.024,
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`Figure 1-7. These sections of a conventional turbofan engine are shown in the
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`cross-section figure below from a 1996 textbook (as well as in the images above of
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`the GE90 engine). The schematic cross section is very similar to the cutaway
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`image but illustrates the major structural components and architecture
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`schematically. A person of ordinary skill in the art would be very familiar with
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`schematic cross section drawings of turbofan engines and would be able to
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`understand major structural components and architecture of an engine from a
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`schematic cross section.
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`10
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`GE-1003.010
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`GE-1014.024, Figure 1-7 (annotations in color)
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`
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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-1014.024-.026; 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 (blue colored) or (2) the core flow path (orange colored). GE-1014.024,
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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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`
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`11
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`GE-1003.011
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`
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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 from approximately 5 to 8.5, meaning that most of the air
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`that enters the engine travels through the bypass flow path. GE-1014.133.
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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,
`
`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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`1014.024-.026, .053-.060. The core flow is first compressed by the low pressure
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`compressor and then the high pressure compressor. GE-1014.024-.026, .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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`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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`12
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`GE-1003.012
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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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`
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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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`1014.024-.026, .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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`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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`
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`13
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`GE-1003.013
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`GE-1014.024, Figure 1-7 (annotations in color)
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`
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`26. Once the core flow is compressed and heated, it must be made to do
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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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`the high spool shaft. GE-1014.024-.026, .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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`14
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`GE-1003.014
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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-1014.024-.026, .057, .088 (“…the low-pressure turbine drives the
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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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`15
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`GE-1003.015
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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-1014.106-.107. The major commercial aircraft propulsion
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`gas turbine manufacturers generally use a conventional two spool configuration2
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`with co-axial high and low spools as shown below:
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`2 Rolls Royce typically utilizes a turbofan configuration that includes three spools.
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`16
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`GE-1003.016
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`GE-1014.024, Figure 1-7 (annotations in color)
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`
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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). As explained in more detail below, each stage of a particular
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`section of the engine (e.g., the low pressure compressor) can be identified by a
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`vertical line that protrudes from the centerline axis, which represents the rotor disc
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`for a stage that is connected to either the low spool shaft or high spool shaft.
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`17
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`GE-1003.017
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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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`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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`18
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`GE-1003.018
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`
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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-1016.014 (“Similar studies on the high-pressure turbine (HPT) disk also
`
`showed an overstressed condition if the bore diameter was increased to
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`accommodate a larger low shaft.”).
`
`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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`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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`
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`19
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`GE-1003.019
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`
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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….”); see also
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`GE-1013.006. Because the fan and low pressure turbine are not directly coupled to
`
`one another, the geared configuration enables the fan to operate at its optimal low
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`rotational speed, while the low pressure turbine operates at its optimal high
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`rotational speed. GE-1013.006 (“The 3:1 speed reduction of the fan drive gear
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`system allows the designer to select the fan tip speed for lowest noise and highest
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`fan efficiency, while at the same time maximizing the low spool shaft speed to
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`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
`
`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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`designed, built, and tested a geared turbofan engine in the 1970s under a contract
`
`from NASA, which was described in a publicly disclosed report (GE-1011):
`
`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
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`
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`20
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`GE-1003.020
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`
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`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. An example of a geared turbofan engine is illustrated below, which
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`comes from a 1990 trade journal that describes development work on ultrahigh
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`bypass geared turbofans performed by Pratt & Whitney (Patent Owner’s
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`subsidiary), MTU, and Fiat Avio:
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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-1013.006 (annotations in red)
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`
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`The geared turbofan shown above4 includes a gear arrangement (“Fan drive gear”)
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`that drives a fan section (“Variable-pitch fan”), a compressor section, a “High-
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`speed low spool” including a three stage low pressure turbine that drives the gear
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`train, and a “slimline nacelle” that forms the bypass airflow path. GE-1013.006.
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`34.
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`It has been well understood in the aviation industry for decades that a
`
`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
`
`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
`
`
`4 The major structures of a turbofan engine are largely symmetric about a center
`
`axis, which is why cross-section images of turbofan engines often depict only the
`
`bottom half or top half of the engine.
`
`
`
`22
`
`GE-1003.022
`
`
`
`
`
`fewer stages when compared to a conventional direct drive turbofan engine, which
`
`can reduce engine weight and maintenance costs. GE-1015.010 (“The [LP
`
`Compressor] is not constrained to rotate at the same speed as the fan”); GE-
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`1013.006 (“The high-speed low spool permits the elimination of a total of three to
`
`five stages cumulatively in the low-pressure compressor and low-pressure
`
`turbine.”).
`
`35.
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`It is well known in the art that a geared configuration can enable an
`
`engine to have a high bypass ratio, which can yield improvements in noise levels
`
`and fuel efficiency. See, e.g., GE-1013.005 (“Since 1985, an extensive technology
`
`readiness program has been under way to demonstrate the key technologies of a
`
`geared, variable-pitch, lightweight fan. and a thin-lip, slim-line nacelle that are
`
`required to make a 10:1 to 20:1 bypass ratio advanced ducted engine a reality.”). It
`
`is also well known that high bypass ratio engines are typically optimized at low fan
`
`pressure ratio5. GE-1013.006 (“Ultrahigh bypass cycles optimize at relatively low
`
`fan-pressure ratios. Depending on the bypass ratio, fan-pressure ratios vary
`
`between 1.2 and 1.4.”); see also GE-1014.077 (Figure 5-29b illustrating that higher
`
`bypass ratio fans optimize at lower fan pressure ratios).
`
`
`5 Fan pressure ratio is typically defined as the ratio of the pressure at the outlet of
`
`the fan section to the pressure at the inlet of the fan section.
`
`
`
`23
`
`GE-1003.023
`
`
`
`
`
`
`
`C.
`
`Fan Section
`
`36. The 412 Patent is directed to attributes of the fan section of a geared
`
`turbofan engine, and more specifically to the number of fan blades (N), the solidity
`
`of the blades at the tip (R), and the ratio of the number of fan blades to the solidity
`
`(N/R).
`
`37. A conventional turbofan engine includes a fan having a number of
`
`blades circumferentially spaced around a hub. GE-1014.025-.028 (illustrating
`
`several turbofan engines). The figure below illustrates a fan section, with each
`
`blade extending radially outward from the hub between a root and a tip. Each
`
`blade also extends in a chord direction between a leading edge and a trailing edge.
`
`The chord dimension can vary over the span of the blade, i.e., the chord dimension
`
`at the root (CDroot) is different than the chord dimension at the tip (CDtip).
`
`
`
`24
`
`
`
`GE-1003.024
`
`
`
`
`
`38. The circumferential pitch of the blades generally refers to the spacing
`
`between the blades. The circumferential pitch of the blades at the tip (CP) is equal
`
`to the circumference of the fan divided by the number of fan blades.6 See, e.g.,
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`GE-1007 at 2:20-22 (“blade pitch is the circumferential length at a given diameter
`
`divided by the number of blades in the full fan blade row.”).
`
`CP(cid:3404) π ∗ Fan Diameter
`Number of Blades (cid:4666)N(cid:4667)
`
`The ratio of the chord dimension (CD) to the circumferential pitch (CP) is known
`
`in the art as solidity. GE-1014.113 (“The ratio of the airfoil chord c to the airfoil
`
`spacing s is called the solidity”).
`
`Solidity (cid:4666)R(cid:4667)(cid:3404)
`
`chord dimension (cid:4666)CD(cid:4667)
`circumferential pitch (cid:4666)CP(cid:4667)
`
`Because the chord dimension (CD) and circumferential pitch (CP) of the blades
`
`vary from root to tip, as described above, a fan section can have a varying solidity
`
`value from root to tip. GE-1009.018, Table I (disclosing a hub (i.e., root) solidity
`
`of 1.0 and tip solidity of 0.67). The solidity value claimed in the 412 Patent is a tip
`
`
`6 The circumferential pitch, like the chord dimension, varies and can be measured
`
`at any point between the root and tip of the blade. The 412 Patent describes and
`
`claims the circumferential pitch of the blades as measured at the tip. GE-1001 at
`
`3:8-10, 4:58-59.
`
`
`
`25
`
`GE-1003.025
`
`
`
`
`
`solidity because the claims describe measuring both CD and CP at the tip. GE-
`
`1001 at 4:56-59 (“extends…between a leading edge and a trailing edge at the tip
`
`to define a chord dimension (CD), said row of propulsor blades defining a
`
`circumferential pitch (CP) with regard to said tips….”).
`
`39. Solidity is a known design parameter that characterizes how much
`
`area the blades sweep through. Lower solidity means less area is swept by the fan
`
`blades, while higher solidity means more area is swept by the fan blades. For
`
`example, given a fan with a diameter of 100 inches and 20 blades, a lower solidity
`
`means a narrower blade chord (e.g., 0.5 solidity = 7.85 inch blade chord), while a
`
`higher solidity means a wider blade chord (e.g., 1.0 solidity = 15.7 inch blade
`
`chord). Similarly, given a fan diameter of 100 inches and a blade chord of 20
`
`inches, a lower solidity means fewer blades (e.g., 0.51 solidity = 8 blades), while a
`
`higher solidity means more blades (e.g., 1.02 solidity = 16 blades).
`
`40. Adjusting the solidity of the fan is known to have various effects.
`
`Increasing the solidity by increasing the chord dimension, for example, can
`
`improve the stability and efficiency of the fan.7 See e.g., GE-1021 at 1:30-32
`
`(“wider chord blades offer the increased efficiency because they have greater
`
`stability margins and move the air more efficiently….”). Increasing the blade
`
`7 A person of ordinary skill would understand, however, that there are limitations
`
`with regards to the efficiency benefits of increasing the blade chord dimension.
`
`
`
`26
`
`GE-1003.026
`
`
`
`
`
`chord dimension, however, also increases the size and weight of the fan. GE-1021
`
`at 1:43-46 (“manufacture of solid titanium wide chord fan blades is prohibitive
`
`because of the…ultimate weight of the blade….”). Conversely, decreasing the
`
`solidity of the fan by either reducing the blade chord dimension or number of
`
`blades reduces the size and weight of the fan, but also reduces the efficiency and
`
`stability of the fan.
`
`41. The 412 Patent also claims a ratio of the number of blades (N) to the
`
`solidity (R). As shown below, the ratio of the number of blades to the solidity
`
`(N/R) is equivalent to the ratio of the circumference of the fan to the blade chord:
`
`(cid:1788)(cid:1792) (cid:3404)N(cid:3400)1R
`(cid:3404) N(cid:3400)CPCD
`(cid:3404)N(cid:3400)π∗Fan Diameter
`N
`(cid:3404)(cid:1780)(cid:1801)(cid:1814) (cid:1777)(cid:1809)(cid:1818)(cid:1803)(cid:1821)(cid:1813)(cid:1806)(cid:1805)(cid:1818)(cid:1805)(cid:1814)(cid:1803)(cid:1805)
`(cid:1777)(cid:1778)
`
`(cid:3400) 1CD
`
`
`
`The 412 Patent explains that the ratio of N/R can be between 8 and 28. GE-1001
`
`at 3:37-40. Put differently, this means that the fan circumference can be between 8
`
`to 28 times the blade chord dimension measured at the tip.
`
`42. The 412 Patent appears to suggest that the following concepts were
`
`new: (i) the disclosed N/R ratios; and (ii) a gear assembly that allows for a low
`
`
`
`27
`
`GE-1003.027
`
`
`
`
`
`speed, large diameter fan. GE-1001 at 4:9-21. As previously explained in ¶¶ 29-
`
`31 above, it was known decades before the filing date of the 412 Patent that a
`
`geared configuration allows for a low speed, large diameter fan. In addition, there
`
`are examples of publications t