# US 10,502,067 B2 · Dual-fed airfoil tip > Agents: this data is available as JSON from the patentgrep API and MCP server — see https://patentgrep.com/llms.txt - Number: US 10,502,067 B2 - Document ID: US10502067B2 - Status: Granted - Kind: B2 - Application number: 15/143,717 - Filed: 2016-05-02 - Published: 2017-07-27 - Granted: 2019-12-10 - Earliest claimed: 2016-01-22 - Assignees: United Technologies Corporation - Applicants: United Technologies Corporation - Inventors: Dominic J. Mongillo; Parth Jariwala; Bret M. Teller; Mark F. Zelesky; James Tilsley Auxier - Web page: https://patentgrep.com/patent/US10502067B2 ## CPC - F01D 5/187: {Convection cooling} - F01D 9/02: Nozzles; Nozzle boxes; Stator blades; Guide conduits {, e.g. individual nozzles (nozzle boxes F01D9/047)} - F01D 25/12: Cooling - F04D 29/324: {Blades} - F04D 29/542: {Bladed diffusers (fixing blades to stators F01D9/042)} - F04D 29/582: {specially adapted for elastic fluid pumps} - F05D 2220/32: in gas turbines - F05D 2250/185: serpentine-like - F05D 2260/202: by film cooling - F05D 2260/22141: using fins or ribs - F05D 2260/606: Bypassing the fluid - Y02T 50/60: Efficient propulsion technologies, e.g. for aircraft - F01D: NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES (machines or engines for liquids F03; non-positive displacement pumps F04D) - F04D: NON-POSITIVE-DISPLACEMENT PUMPS (engine fuel-injection pumps F02M; ion pumps H01J41/12; electrodynamic pumps H02K44/02) - F05D: INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS - Y02T: CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION ## Abstract An airfoil of a gas turbine engine is provided including a leading edge extending in a radial direction, a tip extending in an axial direction from the leading edge, a first rib extending radially within the airfoil, the leading edge and the first rib defining a leading edge cavity within the airfoil, a second rib, the second rib and the first rib defining a serpentine cavity therein, a third rib extending axially within the tip, a flag tip cavity defined by the third rib, the leading edge, and the tip, the leading edge cavity fluidly connected to the flag tip cavity, and a bypass aperture formed between the first rib and the third rib, the bypass aperture configured to fluidly connect the serpentine cavity with the flag tip cavity. ## Claims ### Claim 1 An airfoil of a gas turbine engine comprising: a leading edge extending in a radial direction; a tip extending in an axial direction from the leading edge; a first rib extending radially within the airfoil, the leading edge and the first rib defining a leading edge cavity within the airfoil; a second rib, the second rib and the first rib defining a serpentine cavity therein; a third rib extending axially within the tip, a flag tip cavity defined by the third rib, the leading edge, and the tip, the leading edge cavity fluidly connected to the flag tip cavity; a bypass aperture formed between the first rib and the third rib, the bypass aperture configured to fluidly connect the serpentine cavity with the flag tip cavity, wherein the bypass aperture is located adjacent the first rib; and a divider portion located proximate to the bypass aperture within the serpentine cavity and configured to aid in directing (i) a first portion of air from the serpentine cavity into the flag tip cavity and (ii) a second portion of air within the serpentine cavity, wherein the first rib includes a direction portion proximate to the bypass aperture that is canted toward the leading edge and away from the bypass aperture and divider portion. ### Claim 2 (depends on 1) The airfoil of claim 1, wherein the divider portion is connected to the third rib. ### Claim 3 (depends on 1) The airfoil of claim 1, wherein the divider portion extends a predetermined length into the serpentine cavity. ### Claim 4 (depends on 1) The airfoil of claim 1, wherein the divider portion has a tapered tip extending into the serpentine cavity. ### Claim 5 (depends on 1) The airfoil of claim 1, further comprising a trip strip configured to aid in airflow from the serpentine cavity to the flag tip cavity through the bypass aperture. ### Claim 6 (depends on 1) The airfoil of claim 1, wherein a width of the serpentine cavity between the first rib and the second rib is a first width and a width of the bypass aperture is a second width, wherein the second width is less than the first width. ### Claim 7 (depends on 1) The airfoil of claim 1, further comprising a convective feature within at least one of the leading edge cavity, the flag tip cavity, or the serpentine cavity. ### Claim 8 (depends on 7) The airfoil of claim 7, wherein the convective feature is a plurality of trip strips. ### Claim 9 (depends on 1) The airfoil of claim 1, further comprising film cooling holes configured in at least one of the leading edge, the tip, or the flag tip cavity. ### Claim 10 A gas turbine engine comprising: an airfoil having: a leading edge extending in a radial direction; a tip extending in an axial direction form the leading edge; a first rib extending radially within the airfoil, the leading edge and the first rib defining a leading edge cavity within the airfoil; a second rib, the second rib and the first rib defining a serpentine cavity therein; a third rib extending axially within the tip, a flag tip cavity defined by the third rib, the leading edge, and the tip, the leading edge cavity fluidly connected to the flag tip cavity; a bypass aperture formed between the first rib and the third rib, the bypass aperture configured to fluidly connect the serpentine cavity with the flag tip cavity, wherein the bypass aperture is located adjacent the first rib; and a divider portion located proximate to the bypass aperture within the serpentine cavity and configured to aid in directing (i) a first portion of air from the serpentine cavity into the flag tip cavity and (ii) a second portion of air within the serpentine cavity, wherein the first rib includes a direction portion proximate to the bypass aperture that is canted toward the leading edge and away from the bypass aperture and divider portion. ### Claim 11 (depends on 10) The gas turbine engine of claim 10, wherein the divider portion is connected to the third rib. ### Claim 12 (depends on 10) The gas turbine engine of claim 10, wherein the divider portion extends a predetermined length into the serpentine cavity. ### Claim 13 (depends on 10) The gas turbine engine of claim 10, wherein the divider portion has a tapered tip extending into the serpentine cavity. ### Claim 14 (depends on 10) The gas turbine engine of claim 10, wherein the divider portion is radially angled relative to the axial direction of the third rib. ### Claim 15 (depends on 10) The gas turbine engine of claim 10, wherein a width of the serpentine cavity between the first rib and the second rib is a first width and a width of the bypass aperture is a second width, wherein the second width is less than the first width.