# US 9,752,447 B2 · Angled rail holes > Agents: this data is available as JSON from the patentgrep API and MCP server — see https://patentgrep.com/llms.txt - Number: US 9,752,447 B2 - Document ID: US9752447B2 - Status: Granted - Kind: B2 - Application number: 14/656,446 - Filed: 2015-03-12 - Published: 2016-06-23 - Granted: 2017-09-05 - Earliest claimed: 2014-04-04 - Assignees: United Technologies Corporation - Applicants: United Technologies Corporation - Inventors: Carey Clum; Jonathan Perry Sandoval - Web page: https://patentgrep.com/patent/US9752447B2 ## CPC - F01D 9/023: {Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings} - F01D 5/081: {Cooling fluid being directed on the side of the rotor disc or at the roots of the blades (F01D5/087 takes precedence)} - F02C 7/18: the medium being gaseous, e.g. air {(F02C7/125 takes precedence)} - F23R 3/06: Arrangement of apertures along the flame tube - F05D 2240/81: Cooled platforms - 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) - 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) - F23R: GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS (fluidised bed combustion apparatus specially adapted for operation at superatmospheric pressures F23C10/16) - 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 A gas turbine engine combustor configured to cool a first vane platform by directing airflow through an aft combustor rail is disclosed. In various embodiments, a gas turbine engine combustor may comprise an aft combustor rail connected between an outer wall of an outer liner and an aft heat shield panel, and a plurality of holes in the aft combustor rail. The aft combustor rail may be positioned near an outer vane platform of a first vane stage of a turbine. Further, the plurality of holes may be designed to focus pass-through airflow onto the outer vane platform. In various embodiments, each hole of the plurality of holes may be angled between 20°-90° relative to the aft combustor rail. Additionally, the plurality of holes may be angled to converge the airflow to a focused area, such as a leading edge vane airfoil of the first vane stage. ## Claims ### Claim 1 A gas turbine engine combustor comprising: an aft combustor rail connected between an outer wall of an outer liner and an aft heat shield panel, wherein the aft combustor rail is positioned near an outer vane platform of a first vane stage of a turbine; and a plurality of holes in the aft combustor rail, wherein the plurality of holes is configured to focus pass-through airflow onto the outer vane platform, wherein each hole of the plurality of holes is angled between 20° and 90° in a circumferential direction relative to the aft combustor rail to converge the pass-through airflow to a focused area, wherein the focused area is a leading edge of a vane airfoil of the first vane stage. ### Claim 2 (depends on 1) The gas turbine engine combustor of claim 1, wherein the plurality of holes are parallel to the aft heat shield panel. ### Claim 3 (depends on 1) The gas turbine engine combustor of claim 1, wherein the plurality of holes is angled non-parallel to the aft heat shield panel. ### Claim 4 (depends on 1) The gas turbine engine combustor of claim 1, wherein each hole of the plurality of holes is circular-shaped. ### Claim 5 (depends on 1) The gas turbine engine combustor of claim 1, wherein one or more holes of the plurality of holes is elliptical-shaped. ### Claim 6 (depends on 5) The gas turbine engine combustor of claim 5, wherein one or more holes of the plurality of holes has a diameter in a range of approximately 1 mil to 100 mils. ### Claim 7 (depends on 5) The gas turbine engine combustor of claim 5, wherein one or more holes of the plurality of holes has a length-to-diameter ratio of 1:10. ### Claim 8 (depends on 5) The gas turbine engine combustor of claim 5, wherein the pass-through airflow through the plurality of holes is diverted from an outer annular chamber of the gas turbine engine combustor. ### Claim 9 A gas turbine engine combustor comprising: an aft combustor rail connected between an outer wall of an outer liner and an aft heat shield panel, wherein the aft combustor rail is positioned near an outer vane platform of a first vane stage of a turbine; a first array of circumferentially displaced holes formed through the aft combustor rail, wherein the first array of holes is circumferentially offset in a first circumferential direction from a leading edge of a turbine vane of the first outer vane platform and the first array of holes is angled in a second circumferential direction, the second circumferential direction opposed to the first circumferential direction, to direct pass-through airflow towards the leading edge; and a second array of circumferentially displaced holes formed through the aft combustor rail, wherein second array of holes is circumferentially offset in the second circumferential direction from the leading edge and angled in the first circumferential direction to direct pass-through airflow towards the leading edge. ### Claim 10 (depends on 9) The gas turbine engine combustor of claim 9, wherein the first array of circumferentially displaced holes and the second array of circumferentially displaced holes direct the pass-through airflow to a focused area. ### Claim 11 (depends on 10) The gas turbine engine combustor of claim 10, wherein the focused area of the first array of circumferentially displaced holes is a pressure side of the turbine vane. ### Claim 12 (depends on 9) The gas turbine engine combustor of claim 9, wherein a spacing of multiple holes of the second array of circumferentially displaced holes is greater than a spacing of the multiple holes of the first array of circumferentially displaced holes. ### Claim 13 (depends on 9) The gas turbine engine combustor of claim 9, wherein each hole of the first array of circumferentially displaced holes is angled between 20°-90° relative to the aft combustor rail. ### Claim 14 (depends on 9) The gas turbine engine combustor of claim 9, wherein one or more holes of the first array of circumferentially displaced holes has a diameter in a range of approximately 1 mil to 100 mils. ### Claim 15 (depends on 9) The gas turbine engine combustor of claim 9, wherein one or more holes of the first array of circumferentially displaced holes has a length-to-diameter ratio of 1:10. ### Claim 16 (depends on 9) The gas turbine engine combustor of claim 9, wherein a spacing of multiple holes of the second array of circumferentially displaced holes is greater than a spacing of the multiple holes of the first array of circumferentially displaced holes, wherein each hole of the first array of circumferentially displaced holes is angled between 20° and 90° relative to the aft combustor rail, wherein the angle and the spacing of holes of at least one of the first array of circumferentially displaced holes and the second-array of circumferentially displaced holes creates swirling airflow about the vane airfoil.