# US 2017/0284202 A1 · Fan blade removal feature 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 2017/0284202 A1 - Document ID: US20170284202A1 - Status: Published application - Kind: A1 - Application number: 15/091,066 - Filed: 2016-04-05 - Published: 2017-10-05 - Earliest claimed: 2016-04-05 - Assignees: United Technologies Corporation - Applicants: United Technologies Corporation - Inventors: David A. Welch - Web page: https://patentgrep.com/patent/US20170284202A1 ## CPC - F02K 3/06: with front fan - F01D 5/3007: {of axial insertion type} - F02C 7/04: Air intakes for gas-turbine plants or jet-propulsion plants - F04D 29/322: {Blade mountings} - F04D 29/644: {especially adapted for elastic fluid pumps} - F01D 5/3023: {of radial insertion type, e.g. in individual recesses} - F01D 5/3053: {by means of pins} - F01D 5/323: {Locking of axial insertion type blades by means of a key or the like parallel to the axis of the rotor} - F01D 5/326: {Locking of axial insertion type blades by other means} - F05D 2220/36: specially adapted for the fan of turbofan engines - F05D 2230/70: Disassembly methods - 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) - 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) - 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 ## Abstract A gas turbine engine includes a fan forward of a primary flowpath inlet. The fan includes multiple fan blades distributed radially about, and connected to, a hub. A fan nacelle is positioned radially outward of the fan and includes an inner diameter. The inner diameter is sloped relative to an engine axis such that a forward portion of the fan nacelle has a smaller inner diameter than an aft portion of the fan nacelle. A fan blade spacer is disposed between a radially inward facing surface of at least one fan blade and a radially outward facing surface of the hub. The fan blade spacer has a radial thickness at least equal to a slope drop of the fan nacelle. ## Claims ### Claim 1 A gas turbine engine comprising: a fan forward of a primary flowpath inlet; the fan including a plurality of fan blades distributed radially about, and connected to, a hub; a fan nacelle positioned radially outward of the fan and including an inner diameter, the inner diameter being sloped relative to an engine axis such that a forward portion of the fan nacelle has a smaller inner diameter than an aft portion of the fan nacelle; and a fan blade spacer disposed between a radially inward facing surface of at least one fan blade and a radially outward facing surface of said hub, the fan blade spacer having a radial thickness at least equal to a slope drop of the fan nacelle. ### Claim 2 (depends on 1) The gas turbine engine of claim 1, wherein the spacer has a first radial thickness at a fore edge of the spacer and a second radial thickness at an aft edge of the spacer, and wherein the first radial thickness is distinct from the second radial thickness. ### Claim 3 (depends on 2) The gas turbine engine of claim 2, wherein the first radial thickness is greater than the second radial thickness. ### Claim 4 (depends on 1) The gas turbine engine of claim 1, wherein the radially outward surface of the hub has a first radius at a fore edge of the hub, a second radius at the aft edge of the hub, and wherein the first radius is less than the second radius. ### Claim 5 (depends on 4) The gas turbine engine of claim 4, wherein the fan blade spacer includes a conical center hole. ### Claim 6 (depends on 4) The gas turbine engine of claim 4, wherein a slope of a radially outward facing surface of the hub is greater than or equal to a slope of a radially inward facing surface of a fan nacelle. ### Claim 7 (depends on 6) The gas turbine engine of claim 6, wherein the slope of the radially outward facing surface of the hub is approximately equal to the slope of the radially inward facing surface of the fan nacelle. ### Claim 8 (depends on 1) The gas turbine engine of claim 1, wherein the fan blade spacer includes a center hole. ### Claim 9 (depends on 1) The gas turbine engine of claim 1, wherein the blade spacer is a segmented ring. ### Claim 10 (depends on 1) The gas turbine engine of claim 1, wherein the blade spacer is a single component. ### Claim 11 A method for removing a fan blade from an aircraft engine comprising: at least partially removing a fan blade spacer; shifting a fan blade radially inward, at least a distance equal to a slope drop of a fan nacelle; and sliding said fan blade axially forward. ### Claim 12 (depends on 11) The method of claim 11, wherein shifting the fan blade radially inward further includes shifting a forward edge of the fan blade further radially inward than an aft edge of the fan blade. ### Claim 13 (depends on 12) The method of claim 12, wherein shifting the fan blade further includes shifting the fan blade such that a slope of a fan blade tip is at least a slope of an inner diameter of the fan nacelle. ### Claim 14 (depends on 13) The method of claim 13, wherein shifting the fan blade further includes shifting the fan blade such that a slope of a fan blade tip is approximately equal to a slope of an inner diameter of the fan nacelle. ### Claim 15 (depends on 11) The method of claim 11, wherein removing the fan blade spacer further includes removing a fan blade lock. ### Claim 16 (depends on 11) The method of claim 11, wherein the slope drop of the fan nacelle is defined as a radial increase of an inner diameter of the fan nacelle from an inlet to a leading edge of at least one fan blade. ### Claim 17 (depends on 11) The method of claim 11, wherein the at least partially removing a fan blade spacer comprises removing a segment of the fan blade spacer, the removed segment being radially inward of the fan blade. ### Claim 18 (depends on 11) The method of claim 11, wherein the at least partially removing a fan blade spacer comprises removing a full ring fan blade spacer. ### Claim 19 A fan blade spacer for a gas turbine engine comprising: a ring shaped body including a center hole, the center hole protruding axially through the ring shaped body. ### Claim 20 (depends on 19) The fan blade spacer of claim 19, wherein the center hole is one of a cylindrical hole having a constant diameter and a conical hole having an increasing diameter from a first axial end of the ring shaped body to a second axial end of the ring shaped body opposite the first axial end of the ring shaped body.