# US 10,502,066 B2 · Turbine engine component including an axially aligned skin core passage interrupted by a pedestal > Agents: this data is available as JSON from the patentgrep API and MCP server — see https://patentgrep.com/llms.txt - Number: US 10,502,066 B2 - Document ID: US10502066B2 - Status: Granted - Kind: B2 - Application number: 14/707,118 - Filed: 2015-05-08 - Published: 2016-11-10 - Granted: 2019-12-10 - Earliest claimed: 2015-05-08 - Assignees: United Technologies Corporation - Applicants: United Technologies Corporation - Inventors: Scott D. Lewis - Web page: https://patentgrep.com/patent/US10502066B2 ## CPC - F01D 5/187: {Convection cooling} - F01D 5/147: {Construction, i.e. structural features, e.g. of weight-saving hollow blades (F01D5/148, F01D5/16 and F01D5/20 take precedence; blade shape F01D5/141; blades with cooling or heating channels or cavities F01D5/18; heating, heat-insulating or cooling means on blades F01D5/18)} - F01D 9/041: {using blades (F01D5/148 takes precedence)} - F01D 11/08: for sealing space between rotor blade tips and stator (specially-shaped blade tips therefor F01D5/20) - F01D 25/12: Cooling - F02C 3/04: having a turbine driving a compressor (power transmission arrangements F02C7/36; control of working fluid flow F02C9/16) - F02C 7/18: the medium being gaseous, e.g. air {(F02C7/125 takes precedence)} - F23R 3/002: {Wall structures (F23R3/02 and F23R3/007 take precedence)} - F05D 2220/32: in gas turbines - F05D 2240/127: Vortex generators, turbulators, or the like, for mixing - F05D 2240/30: Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor - F05D 2240/35: Combustors or associated equipment - F05D 2260/202: by film cooling - F05D 2260/2212: by creating turbulence - F05D 2260/2214: by increasing the heat transfer surface - 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 component for a turbine engine includes a fore edge connected to an aft edge via a first surface and a second surface. Multiple cooling passages are defined within the turbine engine component. A skin core passage is defined immediately adjacent the first surface, and at least one pedestal interrupts a flow path through the skin core passage. ## Claims ### Claim 1 A turbine engine component comprising: a fore edge connected to an aft edge via a first surface and a second surface; a plurality of cooling passages defined within the turbine engine component; a skin core passage defined within an exterior wall of the turbine engine component and immediately adjacent said first surface; and at least one pedestal interrupting a flow path through said skin core passage; wherein said at least one pedestal comprises a plurality of pedestals distributed through said skin core passage; wherein said at least one pedestal includes a film cooling hole defined at least partially within the at least one pedestal, the film cooling hole connecting one of said plurality of cooling passages to said first surface; and wherein said film cooling hole connects a first cooling air flow path internal to said turbine engine component to said first surface, and wherein said skin core passage receives cooling air from a second cooling air flow path internal to said turbine engine component, and wherein said second cooling air flow path is a low pressure cooling air flow path relative to said first cooling air flow path. ### Claim 2 (depends on 1) The turbine engine component of claim 1, wherein said at least one pedestal has an oblong cross section relative to a radius of a turbine engine including the turbine engine component. ### Claim 3 (depends on 2) The turbine engine of claim 2, wherein each of said oblong cross sections includes a longest diameter, and said longest diameter is aligned with an axis of the turbine engine. ### Claim 4 (depends on 2) The turbine engine of claim 2, wherein each of said oblong cross sections includes a longest diameter, and said longest diameter is aligned with a radius of the turbine engine. ### Claim 5 (depends on 1) The turbine engine of claim 1, wherein each of said pedestals has a cross section along a direction of flow through the skin core passage, the cross section being one of an oval, a triangle, a rhombus, an airfoil, and a circle. ### Claim 6 (depends on 1) The turbine engine of claim 1, wherein at least 80% of the cooling air entering said skin core passage exits said skin core passage at the aft edge. ### Claim 7 (depends on 6) The turbine engine of claim 6, wherein the skin core passage defines an axial coolant flow path. ### Claim 8 (depends on 1) The turbine engine of claim 1, wherein the turbine engine component is one of a blade outer air seal, a combustor liner, a blade, and a vane. ### Claim 9 (depends on 8) The turbine engine component of claim 8, wherein the blade is a blade in a second or later turbine stage. ### Claim 10 A gas turbine engine comprising: a compressor section; a combustor section fluidly connected to the compressor section by a flowpath; a turbine section fluidly connected to the combustor section by the flowpath; at least one gas turbine engine component exposed to a fluid passing through said flowpath, the at least one gas turbine engine component including: a first surface and a second surface; at least one cooling passage defined within the turbine engine component; a skin core passage defined within an exterior wall of the at least one gas turbine engine component and immediately adjacent said first surface; and at least one pedestal interrupting a flow path through said skin core passage, the at least one pedestal connecting an interior surface of said at least one cooling passage to said first surface; wherein the at least one cooling passage includes a first cooling passage and a second cooling passage, and the at least one pedestal connects the interior surface of the first cooling passage to said first surface; and wherein the first cooling passage receives a first cooling air from a first source, the second cooling passage receives a second cooling air from a second source, and wherein the skin core passage receives the second cooling air from the second cooling passage. ### Claim 11 (depends on 10) The gas turbine engine of claim 10, further comprising a film cooling hole defined within said at least one pedestal, the film cooling hole providing a flow path between said at least one cooling passage to said first surface. ### Claim 12 (depends on 6) The turbine engine of claim 6, wherein at least 95% of the cooling air entering the skin core passage exits the skin core passage at the aft edge. ### Claim 13 (depends on 1) The turbine engine of claim 1, wherein the skin core passage conforms with a contour of the first surface. ### Claim 14 (depends on 13) The turbine engine of claim 13, wherein each of an inward facing wall and an outward facing wall of the skin core passage conforms with the contour of the first surface. ### Claim 15 (depends on 1) The turbine engine of claim 1, wherein the skin core passage is a convective cooling passage.