Ex parte Hasel, Appeal 2017-009971

• Decision Date: 28 November 2018
• Docket Number: Application 13/716,Appeal 2017-009971,253
• Parties: Ex Parte KARL L. HASEL Technology Center 3700
• Court: Patent Trial and Appeal Board

Ex Parte KARL L. HASEL Technology Center 3700

Appeal 2017-009971

Application 13/716, 253

United States Patent and Trademark Office, Patent Trial and Appeal Board

November 28, 2018

FILING DATE: 12/17/2012

Before MICHAEL L. HOELTER, LISA M. GUIJT, and PAUL J. KORNICZKY Administrative Patent Judges.

DECISION ON APPEAL

GUIJT Administrative Patent Judge.

Appellant^[1] appeals under 35 U.S.C § 134(a) from the Examiner's rejection^[2] of claims 1 and 3-14. We have jurisdiction under 35 U.S.C. § 6(b).

We AFFIRM, and we designate our affirmance of the Examiner's rejection of claims 1 and 3-14, under 35 U.S.C. § 112, first paragraph, for failing to comply with the enablement requirement, as a NEW GROUNDS OF REJECTION, pursuant to our authority under 37 C.F.R. § 41.50(b). We enter a NEW GROUNDS OF REJECTION of claims 1 and 3-14 under 35 U.S.C. § 112, first paragraph, for failing to comply with the written description requirement, pursuant to our authority under 37 C.F.R. § 41.50(b).

STATEMENT OF THE CASE

Claims 1 and 11 are the independent claims on appeal. Claim 1, reproduced below, is exemplary of the subject matter on appeal, with disputed claim limitations italicized for emphasis.

1. A gas turbine engine comprising:

a fan rotor, a first compressor rotor and a second compressor rotor, a first turbine rotor and a second turbine rotor, said first compressor rotor configured for operating at a lower pressure than said second compressor rotor and said second turbine rotor configured for operating at a higher pressure than said first turbine rotor, said first turbine rotor configured to drive said first compressor rotor, and said second turbine rotor configured to drive said second compressor rotor and said first turbine rotor also configured to drive said fan rotor through a gear reduction;

wherein a first number is defined as the total of blades collectively associated with each of said fan rotor, said first and second compressor rotors and said first and second turbine rotors,

wherein a second number is defined as the total of static vane members collectively associated with each of said fan rotor, said first and second compressor rotors and said first and second turbine rotors,

wherein a third number is defined as a sum of the first number and the second number;

wherein a fourth number is defined as the total of stages collectively associated with each of the fan rotor, the first and

second compressor rotors and the first and second turbine rotors;

wherein an overall pressure ratio from an inlet end of said fan rotor to an outlet end of said second compressor rotor is configured to be greater than 30 at 35, 000 feet and operating at .80 MN cruise flight condition;

wherein said fan rotor is configured to deliver air into: said first compressor rotor; and a bypass duct as bypass propulsion air,

wherein a bypass ratio is defined as the quantity of air delivered into the bypass duct divided by the quantity of air delivered into the first compressor rotor,

wherein the bypass ratio is greater than about 8.0; wherein a product is defined by the bypass ratio multiplied by the overall pressure ratio, and wherein a stage ratio is defined as said product divided by said fourth number; wherein an airfoil ratio is defined as said product divided by said third number; and

wherein:

said airfoil ratio is greater than or equal to .12; or

said stage ratio is greater than or equal to 22.

THE REJECTIONS

I. Claims l and 3-14 stand rejected under 35U.S.C. § 112, first paragraph, for failing to comply with the enablement requirement.

II. Claims 1 and 3-14 stand rejected under 35 U.S.C. § 112, second paragraph, as being indefinite.

III. Claims 1 and 3-14 stand rejected under 35 U.S.C. § 102(b) as anticipated by Gray (Energy Efficient Engine Preliminary Design and Integration Studies, by D. E. Gray, et al.. Report No. NASA CR-135396, November 1978).

IV. Claims 1 and 3-14 stand rejected under 35 U.S.C. § 103(a), as unpatentable over Gray, Johnson '509 (US 2005/0081509 A1; published Apr. 21, 2005), Daeubler (US 2009/0214824 A1; published Aug. 27, 2009), Senoo (US 2004/0202545 A1; published Oct. 14, 2004), Johnson '650 (US 2008/0141650 A1; published June 19, 2008), Hyde (US 2010/0126178 A1; published May 27, 2010), and Winter (US 2011/0004388 A1; published Jan. 6, 2011).

ANALYSIS

Rejection I

Regarding claims 1, 3, 4, 8, and 10-14, the Examiner determines that the claim recitation of "greater than" is "not bounded at its upper limit," and therefore, such ranges are not enabled by the Specification^[3] because they extend to "infinity." Final Act. 3. In particular, the Examiner determines that "the claimed quantities (bypass ratios, stage ratio, airfoil ratio) do not have an inherent upper limit because the [quantities are] engineering design variable[s]." Id. at 5. See also Ans 3 (determining that the Specification "fails to enable one of ordinary skill in the art [how] to approach the claimed infinite limits."). The Examiner rejects the remaining claims as failing to cure the deficiencies with respect to independent claims 1 and 11. Id. at 5.

Appellant argues that "a claim and specification do not need to 'enable' everything above a claimed boundary.'" Appeal Br. 4. Appellant also argues that "there are clear limits upon the bypass ratio that are inherent in the gas turbine engine art," for example "for the gas turbine engine to work, air must be delivered to the compressor" and "a fan diameter cannot be made so large that the engine would no longer fit in the packaging space on an aircraft." Id. at 4--5. Appellant submits that "a worker of ordinary skill in the art would recognize that a bypass ratio cannot be infinite." Id. at 5. Regarding the other claimed variables with similar ranges. Appellant argues that

[t]he stage ratio and the airfoil ratio are enabled even though there is no upper limit in that a worker of ordinary skill in this art would recognize there is an inherent upper limit. For example, the stage ratio is "said product divided by said fourth number." The "product" is the bypass ratio multiplied by the overall pressure ratio. As previously shown, the bypass ratio cannot be infinite. An overall pressure ratio also cannot be infinite. There are limits on what a compressor section can be called upon to actually do in real-world situations. As the pressure ratio increases, the temperatures and stresses from the pressure increase.

Moreover, the "fourth number" is the total number of stages associated with the engine. For the stage ratio to be infinite, the fourth number would need to approach zero. It would not be possible for a gas turbine engine to operate with no stages in its fan rotor, first and second compressor rotors and first and second turbine rotors. As such a worker of ordinary skill in the art would recognize there is an upper limit to the stage ratio.

The airfoil ratio is defined as the product divided by the third number. The third number is the total number of blades across the fan rotor, first and second compressor rotors and first and second turbine rotors plus the total number of static vanes associated with each of the fan rotor, first and second compressor rotors and first and second turbine rotors. For the airfoil ratio to be infinite, the third number would have to approach zero. A gas turbine engine cannot operate effectively with no blades, or vanes. As such, a worker of ordinary skill in the art would recognize there is an upper limit to the airfoil ratio.

Id. at 5-6; see also Reply Br. 2 (arguing that "limits on the claimed ranges depend on the 'predictable scientific laws' of mechanical arts"). Appellant concludes that "[a] person having skill in the art having read Appellant's disclosure would know how to vary the claimed quantities all the way to their practical limits." Reply Br. 2.

We are persuaded by Appellant's argument that the structural limitations of the claimed gas turbine engine provide inherent upper boundaries on the claimed "greater than" ranges. For example, as argued by Appellant, the fan rotor must be configured to deliver air into a first compressor rotor and a bypass duct, such that the BPR^[4] cannot be infinity, and the claims define structure (i.e., blades, static vane members, and stages) having inherent physical limitations that also provide upper limits to the claimed ranges.

Notwithstanding, we determine that the Specification fails to enable the scope of independent claims l and 11, and claims 3-10 and 12-14 depending therefrom. The statutory basis for the enablement requirement is found in 35 U.S.C. § 112, first paragraph, which provides, in relevant part:

The specification shall contain a written description of the invention, and of the manner and process of making and using it, in such Ml, clear, concise, and exact terms as to enable any person skilled in the art to which it pertains, or with which it is most nearly connected, to make and use the same. . . .

"Enablement serves the dual function in the patent system of ensuring adequate disclosure of the claimed invention and of preventing claims broader than the disclosed invention." MagSil Corp. v. Hitachi Glob. Storage Techs., Inc., 687 F.3d 1377, 1380-81 (Fed. Cir. 2012). When rejecting a claim for lack of enablement, the USPTO bears an initial burden of setting forth a reasonable explanation as to why it believes that the scope of protection provided by that claim is not adequately enabled by the description of the invention provided in the specification. In re Wright, 999 F.2d 1557, 1561-62 (Fed. Cir. 1993). If the USPTO meets this burden, the burden then shifts to the applicant to provide suitable proofs indicating that the specification is indeed enabling. Id. at 1562 (citing In re Marzocchi, 439 F.2d...