Gen. Elec. Co. v. Raytheon Techs. Corp.

• Decision Date: 23 December 2020
• Docket Number: 2019-1319
• Citation: 983 F.3d 1334
• Parties: GENERAL ELECTRIC COMPANY, Appellant v. RAYTHEON TECHNOLOGIES CORPORATION, Appellee
• Court: U.S. Court of Appeals — Federal Circuit

983 F.3d 1334

GENERAL ELECTRIC COMPANY, Appellant

v.RAYTHEON TECHNOLOGIES CORPORATION, Appellee

2019-1319

United States Court of Appeals, Federal Circuit.

Decided: December 23, 2020

William F. Lee, Wilmer Cutler Pickering Hale and Dorr LLP, Boston, MA, argued for appellant. Also represented by Brian Driscoll, Lauren B. Fletcher, Madeleine C. Laupheimer, Louis W. Tompros.

Patrick Joseph Coyne, Finnegan, Henderson, Farabow, Garrett & Dunner, LLP, Washington, DC, argued appellee. Also represented by Sydney Kestle, Jeffrey Curtiss Totten ; Benjamin Aaron Saidman, Atlanta, GA.

Before Lourie, Reyna, and Hughes, Circuit Judges.

Hughes, Circuit Judge.

General Electric Company appeals the Patent Trial and Appeal Board's decision finding Raytheon Technologies Corporation's gas turbine engine patent not unpatentable for obviousness. Raytheon moved to dismiss the appeal for lack of standing. Because General Electric alleged sufficient facts to establish that it is engaging in activity that creates a substantial risk of future infringement, GE has standing to bring its appeal. As to the merits of the appeal, we vacate the Board's decision and remand the case for further consideration because the Board lacked substantial evidence for its conclusions.

I

Raytheon (known as United Technologies Corporation during the appealed proceedings) and GE vigorously compete in the market to supply propulsion engines to the commercial aviation industry. This dispute revolves around the validity of Raytheon's patent's claims to a two-stage high pressure turbine engine for commercial airplanes and whether those claims would have been obvious in light of the prior art.

A

We begin with a brief technical background. This dispute centers on turbofan gas turbine engines used to propel commercial airliners. See J.A. 1182. Turbofan engines rely on four main component sections—the fan, compressor, combustor, and turbine—to generate thrust from the continuous ignition of a mixture of fuel and pressurized air. J.A. 1183.

J.A. 1184. To do so, air enters the fan, which accelerates the air using rotating airfoil "blades. " Id. The specific engines here are high-bypass-ratio turbofans, in which a portion of the air, after passing through the fan, immediately exits the engine to generate thrust from the momentum imparted upon it by the fan. J.A. 1185. That air is known as the "bypass flow." Id. The rest of the air from the fan enters the engine "core," or the compressor, combustor, and turbine sections. Id. That air is known as the "core flow." Id. The ratio of bypass flow to core flow is called the bypass ratio. For commercial airliners, a higher bypass ratio (i.e., more bypass flow for a given amount of core flow) increases fuel efficiency. See J.A. 1797.

The compressor and turbine sections are further divided into high- and low-pressure segments. J.A. 1183. Each of the high- and low-pressure compressor and turbine sections consist of stages, or a "a matched set of rotating blades and stationary airfoils." See Gen. Elec. Co. v. United Techs. Corp. , No. IPR2017-00428, 2018 WL 3105491, at *7 n.6 (P.T.A.B. June 22, 2018) ( Final Written Decision ); J.A. 1186. In the figure above, these stages are represented by vertical black lines extending from the central axis in the compressor and turbine sections of the engine. J.A. 1186, n.1. Core flow air is pressurized in the compressor section before it enters the combustor, where it is mixed with fuel and ignites. The resulting hot gas enters the turbine where the expansion of the gas powers the turbine's rotating blades. See Appellant's Br. 7–8.

Artisans refer to the grouping of the high-pressure compressor and high-pressure turbine as the "high [pressure] spool" and the grouping of the fan, low-pressure compressor, and low-pressure turbine as the "low spool." Id. In a conventional "direct-drive" turbofan engine the components comprising the low spool are all connected to the same shaft and rotate at the same speed. See Appellant's Br. 8. The technology here, however, involves a "geared" turbofan, which uses a gearbox mounted between the low-pressure compressor and the fan to reduce the rotational speed of the fan compared to the low-pressure compressor and low-pressure turbine. Id. In a high-bypass-ratio turbofan, the fan has a much larger diameter than the engine core components; this discrepancy in diameter leads to a discrepancy in the ideal rotational speed of the fan compared to the low-pressure turbine. See J.A. 3269.

By introducing a gearbox that allows the fan to rotate more slowly than the rest of the low spool, each component can run at an operating point much closer to its optimal rotational speed, yielding many benefits. Some of these benefits include:

(1) improving the propulsive efficiency of the fan, reducing engine fuel consumption;

(2) improving the aerodynamic efficiency of the low-pressure turbine, allowing a simpler and less costly design;

(3) reducing the mechanical stress on the fan, improving safety and reliability;

(4) reducing torque on the low-spool shaft connecting the low-pressure compressor and turbine to the fan (or gearbox), allowing the use of a smaller-diameter shaft; and;

(5) reducing engine noise caused by high fan rotational speeds. See, e.g. , J.A. 1333, 1351, 1361, 1797–98.

B

In 2011, Raytheon applied for the patent that issued as U.S. Patent Number 8,695,920, entitled "Gas Turbine Engine with Low Stage Count Low Pressure Turbine." According to the background, the invention relates to "an engine mounting configuration for the mounting of a turbofan gas turbine engine to an aircraft pylon." ’920 patent at 1:13–15. Although much of the written description focuses on the "static structure" of the engine used to help mount the turbofan to an aircraft, see, e.g. , ’920 patent at cols. 5–7, the patent claims certain inventions involving particular turbofan gas turbine engine configurations. The claims in dispute, dependent claims 10–14, relate to a "method of designing a gas turbine engine" comprising certain of these architectural features and performance parameters. Each claim depends from independent claim 9, reproduced below.

9. A method of designing a gas turbine engine comprising:

providing a core nacelle defined about an engine centerline axis;

providing a fan nacelle mounted at least partially around said core nacelle to define a fan bypass flow path for a fan bypass airflow;

providing a gear train within said core nacelle; providing a first spool along said engine centerline axis within said core nacelle to drive said gear train, said first spool includes a first turbine section including between three–six (3–6) stages, and a first compressor section;

providing a second spool along said engine centerline axis within said core nacelle, said second spool includes a second turbine section including at least two (2) stages and a second compressor section;

providing a fan including a plurality of fan blades to be driven through the gear train by the first spool, wherein the bypass flow path is configured to provide a bypass ratio of airflow through the bypass flow path divided by airflow through the core nacelle that is greater than about six (6) during engine operation.

’920 patent at 8:14–34. Claim 10 claims the method of claim 9, adding the limitation that "said first turbine section defines a pressure ratio that is greater than about five (5.0)." Id. at 8:35–37. Claim 11 depends from claim 10's method, adding the limitation that "a fan pressure ratio across the plurality of fan blades is less than about 1.45." Id. at 8:38–41. Claim 12 depends from claim 11, adding the limitation that "the gear train is configured to provide a speed reduction ratio greater than about 2.5:1." Id. at 8:42–44. Claim 13 depends from claim 12, further requiring that "the plurality of fan blades [be] configured to rotate at a fan tip speed of less than about 1150 feet/second during engine operation." Id. at 8:44–46. Finally, claim 14 depends from claim 13, requiring, together with all the limitations disclosed in claims 9–13, that "the second turbine section includes two (2) stages." Id. at 8:47–48.

As the parties do, we will refer to the "first" spool, turbine, and compressor as the low-pressure spool, turbine, and compressor; the "second" spool, turbine, and compressor, as the high-pressure spool, turbine, and compressor.

The ’920 patent issued in April 2014. In December 2016, GE petitioned the Patent Trial and Appeal Board for inter partes review of claims 1–4, 7–14, 17, and 19 of the ’920 patent. In its petition, GE asserted that claims 1, 4, 9–14, 17, and 19 were unpatentable as obvious based on the combination of two prior art references, Wendus and Moxon.

C

Wendus—Follow-On Technology Requirement Study for Advanced Subsonic Transport , by Bruce E. Wendus, et al.—is a research paper by four employees of Raytheon subsidiary Pratt & Whitney, published internally in August 1995 before being publicly distributed as a NASA "contractor report" in August 2003. Wendus provides a computational study comparing a "baseline," "state-of-the-art" 1995 turbofan engine to a hypothetical advanced technology engine it dubs the "Advanced Ducted Propulsor" that incorporates technology thought to be feasible for engines entering service in 2005. J.A. 1310–12. Wendus compares these engines to determine the advanced engine's performance gains and its effects on the economics of airline operation compared to the baseline. Id. Wendus then suggests future research for technologies that are "critical or enabling" to implementing the advanced engine's design and attaining its performance advantages. J.A. 1311. In its description of the advanced engine, Wendus discloses all elements of claims 9–14 except that it teaches a one-stage high-pressure turbine instead of the "at-least-two-stage" high-pressure turbine taught in claim 9 and narrowed to two stages in claim 14. See J.A. 16.

Moxon is a July 1983 magazine article—How to Save Fuel in...