# US 2018/0320633 A1 · Aircraft and Direct Drive Engine Under Wing Installation > Agents: this data is available as JSON from the patentgrep API and MCP server — see https://patentgrep.com/llms.txt - Number: US 2018/0320633 A1 - Document ID: US20180320633A1 - Status: Published application - Kind: A1 - Application number: 15/439,167 - Filed: 2017-02-22 - Published: 2018-11-08 - Earliest claimed: 2017-02-22 - Assignees: General Electric Company - Applicants: General Electric Company - Inventors: Thomas Ory Moniz; Randy M. Vondrell; Jeffrey Donald Clements; Brandon Wayne Miller - Web page: https://patentgrep.com/patent/US20180320633A1 ## CPC - F02K 3/075: controlling flow ratio between flows - B64D 27/12: within, or attached to, wings - B64D 29/02: associated with wings - F01D 21/045: {special arrangements in stators or in rotors dealing with breaking-off of part of rotor} - F02C 3/107: with two or more rotors connected by power transmission - F02C 3/13: having variable working fluid interconnections between turbines or compressors or stages of different rotors {(controlling flow ratio between different flows of multi-flow jet-propulsion plant, e.g. ducted fan F02K3/075)} - F02C 7/20: Mounting or supporting of plant; Accommodating heat expansion or creep - F02C 7/36: Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user ({F02C3/107 - F02C3/13 and} F02C7/32 take precedence; couplings for transmitting rotation F16D; gearing in general F16H) - F02K 3/06: with front fan - B64D 33/04: of exhaust outlets or jet pipes - F05D 2260/40: Transmission of power - F05D 2300/433: Polyamides, e.g. NYLON - F05D 2300/601: Fabrics - F05D 2300/6033: Ceramic matrix composites [CMC] - Y02T 50/40: Weight reduction - 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) - B64D: EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN 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) - 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 The present disclosure is directed to a gas turbine engine defining a radial direction, a longitudinal direction, and a circumferential direction, an upstream end and a downstream end along the longitudinal direction, and an axial centerline extended along the longitudinal direction. The gas turbine engine includes a fan assembly including a plurality of fan blades rotatably coupled to a fan rotor in which the fan blades define a maximum fan diameter and a fan pressure ratio. The gas turbine engine further includes a low pressure (LP) turbine defining a core flowpath therethrough generally along the longitudinal direction. The core flowpath defines a maximum outer flowpath diameter relative to the axial centerline. The gas turbine engine defines a fan to turbine diameter ratio of the maximum fan diameter to the maximum outer flowpath diameter. The fan to turbine diameter ratio over the fan pressure ratio is approximately 0.90 or greater. ## Claims ### Claim 1 A gas turbine engine defining a radial direction, a longitudinal direction, and a circumferential direction, an upstream end and a downstream end along the longitudinal direction, and an axial centerline extended along the longitudinal direction, the gas turbine engine comprising: a fan assembly comprising a plurality of fan blades rotatably coupled to a fan rotor, the fan blades defining a maximum fan diameter and a fan pressure ratio; and a low pressure (LP) turbine defining a core flowpath therethrough generally along the longitudinal direction, wherein the core flowpath defines a maximum outer flowpath diameter relative to the axial centerline, and wherein the gas turbine engine defines a fan to turbine diameter ratio of the maximum fan diameter to the maximum outer flowpath diameter, and wherein the fan to turbine diameter ratio over the fan pressure ratio is approximately 0.90 or greater. ### Claim 2 (depends on 1) The gas turbine engine of claim 1, wherein the fan assembly defines a fan pressure ratio between approximately 1.0 and approximately 1.8. ### Claim 3 (depends on 2) The gas turbine engine of claim 2, wherein the fan to turbine diameter ratio over the fan pressure ratio is approximately 1.20 or greater at a fan pressure ratio of approximately 1.50 or less. ### Claim 4 (depends on 2) The gas turbine engine of claim 2, wherein the fan to turbine diameter ratio over the fan pressure ratio is approximately 1.05 or greater at a fan pressure ratio of between approximately 1.50 to approximately 1.60. ### Claim 5 (depends on 1) The gas turbine engine of claim 1, wherein the fan to turbine diameter ratio over the fan pressure ratio is approximately 0.90 or greater at a fan pressure ratio of between approximately 1.60 and approximately 1.80. ### Claim 6 (depends on 1) The gas turbine engine of claim 1, wherein the fan to turbine diameter ratio is approximately 1.8 or greater. ### Claim 7 (depends on 1) The gas turbine engine of claim 1, wherein the LP turbine and the fan rotor of the fan assembly are rotatably coupled in direct drive configuration via a driveshaft. ### Claim 8 (depends on 1) The gas turbine engine of claim 1, wherein the LP turbine defines six or more stages of rotors. ### Claim 9 (depends on 8) The gas turbine engine of claim 8, wherein the LP turbine defines the maximum outer flowpath diameter at the two or more rotating stages at the downstream-most end of the LP turbine. ### Claim 10 (depends on 9) The gas turbine engine of claim 9, wherein the LP turbine defines the maximum outer flowpath diameter at the three or more rotating stages at the downstream-most end of the LP turbine. ### Claim 11 (depends on 1) The gas turbine engine of claim 1, further comprising a containment shield, wherein the containment shield is coupled to an outer casing of the engine extended generally along the longitudinal direction, wherein the containment shield extends at least partially along the circumferential direction from the top dead center reference line along the clockwise and/or counter-clockwise direction. ### Claim 12 (depends on 11) The gas turbine engine of claim 11, wherein the containment shield extends approximately 60 degrees or less clockwise and/or counter-clockwise from the top dead center reference line along the circumferential direction. ### Claim 13 An aircraft defining a longitudinal direction, a latitudinal direction, and a transverse direction, the aircraft including a fuselage extended along the longitudinal direction to which a pair or more of wings attaches along the transverse direction, the aircraft comprising: a wing extended from the fuselage, the wing comprising a pylon, and wherein the wing defines a leading edge and a trailing edge, and wherein the leading edge defines a forward plane extended along a radial direction, and wherein the trailing edge defines an aft plane extended along the radial direction; and a gas turbine engine coupled to the pylon of the wing, wherein each engine comprises: a fan assembly comprising a plurality of fan blades rotatably coupled to a fan rotor, the fan blades defining a maximum fan diameter and a fan pressure ratio; and a low pressure (LP) turbine defining a core flowpath therethrough generally along the longitudinal direction, wherein the core flowpath defines a maximum outer flowpath diameter relative to the axial centerline, and wherein the gas turbine engine defines a fan to turbine diameter ratio of the maximum fan diameter to the maximum outer flowpath diameter, and wherein the fan to turbine diameter ratio over the fan pressure ratio is approximately 0.90 or greater. ### Claim 14 (depends on 13) The aircraft of claim 13, wherein the LP turbine further defines a first turbine rotor at an upstream-most end of the LP turbine and a last turbine rotor at a downstream-most end of the LP turbine, and further wherein the LP turbine defines a turbine burst area inward of the wing along the radial direction, the turbine burst area extended from the axial centerline at a first angle along a plane of rotation of the first turbine rotor toward the upstream end, and at a second angle along a plane of rotation of the last turbine rotor toward the downstream end. ### Claim 15 (depends on 14) The aircraft of claim 14, wherein the first angle of the turbine burst area is approximately 15 degrees or less. ### Claim 16 (depends on 14) The aircraft of claim 14, wherein the second angle of the turbine burst area is approximately 15 degrees or less. ### Claim 17 (depends on 13) The aircraft of claim 13, wherein the wing defines a wing shear center, and wherein the gas turbine engine further comprises: an exhaust nozzle disposed downstream of the LP turbine, wherein the exhaust nozzle defines a downstream-most end, and wherein the downstream-most end is approximately equal to the wing shear center of the wing along the longitudinal direction. ### Claim 18 (depends on 13) The aircraft of claim 13, wherein the fan assembly of the gas turbine engine defines a fan pressure ratio between approximately 1.0 and approximately 1.8. ### Claim 19 (depends on 13) The aircraft of claim 13, wherein the fan to turbine diameter ratio over the fan pressure ratio is approximately 1.20 or greater at a fan pressure ratio of approximately 1.50 or less. ### Claim 20 (depends on 13) The aircraft of claim 13, wherein the fan to turbine diameter ratio of the gas turbine engine is approximately 1.8 or greater.