# US 2023/0056571 A1 · Front section stiffness ratio > Agents: this data is available as JSON from the patentgrep API and MCP server — see https://patentgrep.com/llms.txt - Number: US 2023/0056571 A1 - Document ID: US20230056571A1 - Status: Published application - Kind: A1 - Application number: 17/512,824 - Filed: 2021-10-28 - Published: 2023-02-23 - Earliest claimed: 2021-08-20 - Assignees: Raytheon Technologies Corporation - Applicants: Raytheon Technologies Corporation - Inventors: Marc J. Muldoon; Michael E. McCune; Keith B. Allyn - Web page: https://patentgrep.com/patent/US20230056571A1 ## CPC - F01D 25/28: Supporting or mounting arrangements, e.g. for turbine casing - F02K 3/06: with front fan - F01D 9/02: Nozzles; Nozzle boxes; Stator blades; Guide conduits {, e.g. individual nozzles (nozzle boxes F01D9/047)} - F01D 25/24: Casings (modified for heating or cooling F01D25/14); Casing parts, e.g. diaphragms, casing fastenings (casings for rotary machines or engines in general F16M {; special arrangements in stators dealing with breaking-off of part of rotor F01D21/045}) - F02C 7/20: Mounting or supporting of plant; Accommodating heat expansion or creep - F04D 29/542: {Bladed diffusers (fixing blades to stators F01D9/042)} - F04D 29/665: {by means of resonance chambers or interference} - F05D 2220/36: specially adapted for the fan of turbofan engines - F05D 2240/12: Fluid guiding means, e.g. vanes - F05D 2240/14: Casings or housings protecting or supporting assemblies within - F05D 2240/90: Mounting on supporting structures or systems - F05D 2250/291: hollowed - F05D 2250/75: given by its similarity to a letter, e.g. T-shaped - F05D 2260/30: Retaining components in desired mutual position - F05D 2260/40311: of the epicyclical, planetary or differential type - F05D 2260/606: Bypassing the fluid - F05D 2260/96: Preventing, counteracting or reducing vibration or noise - F05D 2300/60: Properties or characteristics given to material by treatment or manufacturing - 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) - 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) - 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 - Y02T: CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION ## Abstract A gear reduction reduces a speed of a fan rotor relative to a speed of a fan drive turbine. A rigid connection between a fan case and an inner core housing includes a plurality of A-frames connected at a connection point to the fan case. Legs in the A-frames extend away from the connection point in opposed circumferential directions to be connected to a compressor wall of the inner core housing. The rigid connection also includes a plurality of fan exit guide vanes rigidly connected to the fan case. A lateral stiffness ratio of the lateral stiffness of the plurality of fan exit guide vanes and a lateral stiffness of a combination of the plurality of A-frame, the compressor wall, and a fan intermediate case which is forward of the low pressure compressor being greater than or equal to 0.6 and less than or equal to 2.0. ## Claims ### Claim 1 A gas turbine engine comprising: a fan rotor driven by a fan drive turbine about an axis through a gear reduction to reduce a speed of said fan rotor relative to a speed of said fan drive turbine; a fan case surrounding said fan rotor, and an inner core housing surrounding a compressor section, including a low pressure compressor; said fan rotor delivering air into a bypass duct defined between said fan case and a splitter wall of said inner core housing, and a rigid connection between said fan case and said inner core housing including a plurality of A-frames each including a pair of legs rigidly connected at a connection point to said fan case, and each leg in said pair extending away from said connection point in opposed circumferential directions to be connected to a compressor wall of said inner core housing to form an A-shape; and the rigid connection also including a plurality of fan exit guide vanes rigidly connected to said fan case; and a lateral stiffness is defined in a radial direction, a lateral stiffness ratio of the lateral stiffness of the plurality of fan exit guide vanes and a lateral stiffness of a combination of the plurality of A-frame, the compressor wall, and a fan intermediate case which is forward of the low pressure compressor being greater than or equal to 0.6 and less than or equal to 2.0. ### Claim 2 (depends on 1) The gas turbine engine as set forth in claim 1, wherein a moment stiffness is defined about an axis extending perpendicular to a rotational axis of the gas turbine engine, and a moment stiffness ratio of the moment stiffness of the plurality of fan exit guide vanes to a moment stiffness of the combination of the plurality of A-frames, the compressor wall, and the fan intermediate case being greater than or equal to 0.50 and less than or equal to 2.0. ### Claim 3 (depends on 2) The gas turbine engine as set forth in claim 2, wherein the moment stiffness ratio being less than or equal to 1.4. ### Claim 4 (depends on 2) The gas turbine engine as set forth in claim 2, wherein the moment stiffness ratio being less than or equal to 1.0. ### Claim 5 (depends on 2) The gas turbine engine as set forth in claim 2, wherein a first gap is defined between a leading edge of the fan exit guide vanes at a radially innermost point radially outward of the splitter wall and a leading edge of the plurality of A-frames at a radially inner location radially inward of the splitter wall, and a second gap is defined between a trailing edge of the fan blades and a leading edge of the plurality of fan exit guide vanes at the radially inward point radially outward of the splitter wall each of the gaps being defined by extending a line from the respective points in a radially outward direction, and then measuring a distance with a line extending parallel to the rotational axis of the engine, and a first gap ratio of said second gap to said first gap is between 0.25 and 0.45. ### Claim 6 (depends on 5) The gas turbine engine as set forth in claim 5, wherein a third gap is defined between the trailing edge of the fan blades and a leading edge of the fan exit guide vane at a radially outermost location radially inward of the fan case, the third gap also defined by extending a line from the respective points in a radially outward direction, and then measuring a distance with a line extending parallel to the rotational axis of the engine, and a second gap ratio of said first gap to said third gap is greater than or equal to 1.1 and less than or equal to 2. ### Claim 7 (depends on 6) The gas turbine engine as set forth in claim 6, wherein a third gap ratio of said second gap to said third gap is greater than or equal to 0.4 and less than or equal to 0.6. ### Claim 8 (depends on 2) The gas turbine engine as set forth in claim 2, wherein the lateral stiffness ratio being less than or equal to 1.4. ### Claim 9 (depends on 8) The gas turbine engine as set forth in claim 8, wherein the lateral stiffness ratio being less than or equal to 1.0. ### Claim 10 (depends on 1) The gas turbine engine as set forth in claim 1, wherein a ratio of a quantity of the plurality of fan exit guide vanes to a quantity of the legs of the plurality of A-frames being greater than or equal to 1.75. ### Claim 11 (depends on 1) The gas turbine engine as set forth in claim 1, wherein said low pressure compressor has between four to six stages. ### Claim 12 (depends on 1) The gas turbine engine as set forth in claim 1, wherein a first gap is defined between a leading edge of the fan exit guide vanes at a radially innermost point radially outward of the splitter wall and a leading edge of the plurality of A-frames at a radially inner location radially inward of the splitter wall, and a second gap is defined between a trailing edge of the fan blades and a leading edge of the plurality of fan exit guide vanes at the radially inward point radially outward of the splitter wall each of the gaps being defined by extending a line from the respective points in a radially outward direction, and then measuring a distance with a line extending parallel to the rotational axis of the engine, and a first gap ratio of said second gap to said first gap is between 0.25 and 0.45. ### Claim 13 (depends on 11) The gas turbine engine as set forth in claim 11, wherein a third gap is defined between the trailing edge of the fan blades and a leading edge of the fan exit guide vane at a radially outermost location radially inward of the fan case, the third gap also defined by extending a line from the respective points in a radially outward direction, and then measuring a distance with a line extending parallel to the rotational axis of the engine, and a second gap ratio of said first gap to said third gap is greater than or equal to 1.1 and less than or equal to 2. ### Claim 14 (depends on 13) The gas turbine engine as set forth in claim 13, wherein a third gap ratio of said second gap to said third gap is greater than or equal to 0.4 and less than or equal to 0.6. ### Claim 15 (depends on 1) The gas turbine engine as set forth in claim 1, wherein the lateral stiffness ratio being less than or equal to 1.4. ### Claim 16 (depends on 15) The gas turbine engine as set forth in claim 15, wherein the lateral stiffness ratio being less than or equal to 1.0. ### Claim 17 A gas turbine engine comprising: a fan rotor driven by a fan drive turbine about an axis through a gear reduction to reduce a speed of said fan rotor relative to a speed of said fan drive turbine; a fan case surrounding said fan rotor, and an inner core housing surrounding a compressor section, including a low pressure compressor; said fan rotor delivering air into a bypass duct defined between said fan case and a splitter wall of said inner core housing, and a rigid connection between said fan case and said inner core housing including a plurality of A-frames each including a pair of legs rigidly connected at a connection point to said fan case, and each leg in said pair extending away from said connection point in opposed circumferential directions to be connected to a compressor wall of said inner core housing to form an A-shape; the rigid connection also including a plurality of fan exit guide vanes rigidly connected to said fan case; and wherein a moment stiffness is defined about an axis extending perpendicular to a rotational axis of the gas turbine engine, and a moment stiffness ratio of a moment stiffness of the plurality of fan exit guide vanes to a moment stiffness of the combination of the plurality of A-frames, the compressor wall, and a fan intermediate case forward of the lower pressure compressor being greater than or equal to 0.50 and less than or equal to 2.0. ### Claim 18 (depends on 17) The gas turbine engine as set forth in claim 17, wherein the moment stiffness ratio being less than or equal to 1.4. ### Claim 19 (depends on 18) The gas turbine engine as set forth in claim 18, wherein the moment ratio being less than or equal to 1.0. ### Claim 20 (depends on 17) The gas turbine engine as set forth in claim 17, wherein a ratio of a quantity of the plurality of fan exit guide vanes to a quantity of the legs of the plurality of A-frames being greater than or equal to 1.75.