# US 10,371,092 B2 · Convergent divergent exit nozzle for a gas turbine engine > Agents: this data is available as JSON from the patentgrep API and MCP server — see https://patentgrep.com/llms.txt - Number: US 10,371,092 B2 - Document ID: US10371092B2 - Status: Granted - Kind: B2 - Application number: 14/817,611 - Filed: 2015-08-04 - Published: 2016-07-07 - Granted: 2019-08-06 - Earliest claimed: 2014-08-08 - Assignees: United Technologies Corporation - Applicants: United Technologies Corporation - Inventors: Dmitriy B. Sidelkovskiy; Oleg Petrenko; Robert E. Malecki; Steven H. Zysman - Web page: https://patentgrep.com/patent/US10371092B2 ## CPC - F02K 1/78: Other construction of jet pipes - B64D 29/00: Power-plant nacelles, fairings or cowlings - F02C 3/04: having a turbine driving a compressor (power transmission arrangements F02C7/36; control of working fluid flow F02C9/16) - F02K 1/00: Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto (rocket nozzles F02K9/97) - F02K 1/36: having an ejector - F02K 1/40: Nozzles having means for dividing the jet into a plurality of partial jets or having an elongated cross-section outlet - F02K 1/52: Nozzles specially constructed for positioning adjacent to another nozzle or to a fixed member, e.g. fairing - F02K 1/82: Jet pipe walls, e.g. liners - F02K 3/06: with front fan - F02K 3/077: the plant being of the multiple flow type, i.e. having three or more flows - B64D 29/02: associated with wings - B64D 29/04: associated with fuselages - F02K: JET-PROPULSION PLANTS (features of jet-propulsion plants common to gas-turbine plants, air intakes or fuel supply control of air-breathing jet-propulsion plants F02C7/00, F02C9/00) - B64D: EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT - F02C: GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS (construction of turbines F01D; jet-propulsion plants F02K; construction of compressors or fans F04; generating combustion products of high pressure or high velocity F23R; using gas turbines in compression refrigeration plants F25B11/00) ## Abstract A nacelle for a gas turbine engine includes a ring shaped body defining a center axis and having a radially outward surface and a radially inward surface. An aft portion of the radially inward surface includes an axially extending convergent-divergent exit nozzle. An axially extending secondary duct passes through the nacelle in the convergent-divergent exit nozzle. The axially extending secondary duct includes an inlet at a convergent portion of the convergent-divergent exit nozzle and an outlet at a divergent portion of the convergent-divergent exit nozzle. ## Claims ### Claim 1 A method for controlling flow separation in a convergent-divergent nozzle of a gas turbine engine fan duct comprising: receiving a portion of bypass airflow through a primary fan duct into a secondary passage via a secondary passage inlet in a convergent region of the convergent-divergent nozzle; expelling air from said secondary passage through a secondary passage outlet nozzle, wherein the secondary passage outlet nozzle is positioned at least partially in a divergent region of the convergent-divergent nozzle, thereby altering a flow separation of airflow through the primary fan duct in the divergent region; and wherein the secondary passage includes a circumferential ring duct and a plurality of isolated airflow passages, each of the isolated airflow passages extending a partial length of the secondary passage and being contained entirely within the divergent region of the convergent-divergent nozzle. ### Claim 2 (depends on 1) The method of claim 1, wherein receiving the portion of bypass airflow through said primary fan duct into the secondary passage via the secondary passage inlet in the convergent region of the convergent-divergent nozzle comprises receiving said portion of said bypass airflow into the circumferential ring duct. ### Claim 3 (depends on 2) The method of claim 2, further comprising splitting said received portion of said bypass airflow into the plurality of isolated airflow passages after said received portion of said bypass airflow has passed through said circumferential ring duct. ### Claim 4 A nacelle for a gas turbine engine, the nacelle comprising: a ring shaped structure defining a center axis and having a radially outward surface and a radially inward surface; an aft portion of the radially inward surface including an axially extending convergent-divergent exit nozzle; and an axially extending secondary duct passing through said nacelle in said convergent-divergent exit nozzle, the axially extending secondary duct including an inlet at a convergent portion of the axially extending convergent-divergent exit nozzle and an outlet at a divergent portion of the axially extending convergent-divergent exit nozzle, the axially extending secondary duct including a circumferential ring duct and a plurality of isolated airflow passages, each of the isolated airflow passages extending a partial length of the axially extending secondary duct and contained entirely within the divergent portion of the axially extending convergent-divergent exit nozzle. ### Claim 5 (depends on 4) The nacelle of claim 4, wherein the plurality of isolated airflow passages are aft of the circumferential ring duct, relative to a fluid flow through the axially extending secondary duct. ### Claim 6 (depends on 4) The nacelle of claim 4, further comprising a plurality of structures disposed circumferentially about said center axis within said axially extending secondary duct, and wherein said plurality of structures supports a portion of the axially extending convergent-divergent exit nozzle, the portion being radially inward of the axially extending secondary duct relative to a main nacelle portion. ### Claim 7 A gas turbine engine comprising: an engine core and a fan driven by the engine core; a nacelle radially surrounding the engine core and the fan; a bypass passage including a convergent-divergent exit nozzle, the bypass passage being defined by the nacelle at an outer diameter of the bypass passage and by an outer diameter of the engine core at an inner diameter of the bypass passage, wherein the convergent-divergent exit nozzle includes: a primary passage and a secondary passage radially exterior to the primary passage, wherein the secondary passage includes a secondary passage inlet disposed in a convergent region of the convergent-divergent exit nozzle, a secondary passage outlet disposed in a divergent region of the convergent-divergent exit nozzle, a circumferential ring duct, and a plurality of isolated airflow passages, each of the isolated airflow passages extending a partial length of the secondary passage and being contained entirely within the divergent portion of the convergent-divergent exit nozzle. ### Claim 8 (depends on 7) The gas turbine engine of claim 7, wherein 100% of air passing through the secondary passage originates in the primary passage. ### Claim 9 (depends on 7) The gas turbine engine of claim 7, wherein the circumferential ring duct is an unobstructed circumferential ring duct. ### Claim 10 (depends on 7) The gas turbine engine of claim 7, wherein the plurality of isolated airflow passages are aft of the circumferential ring duct, relative to a fluid flow through the secondary passage, and wherein each of the plurality of isolated airflow passages are directly connected to the secondary passage outlet. ### Claim 11 (depends on 7) The gas turbine engine of claim 7, further comprising a plurality of structures disposed circumferentially within said secondary passage, and wherein said plurality of structures supports a portion of the convergent-divergent exit nozzle, the portion being radially inward of the secondary passage. ### Claim 12 (depends on 7) The gas turbine engine of claim 7, wherein the secondary passage outlet of the secondary passage and an outlet of the convergent-divergent exit nozzle are axially aligned, relative to an axis defined by the gas turbine engine. ### Claim 13 (depends on 7) The gas turbine engine of claim 7, wherein the secondary passage exclusively receives air from the secondary passage inlet.