# US 11,898,518 B2 · Aircraft and direct drive engine under wing installation

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- Number: US 11,898,518 B2
- Document ID: US11898518B2
- Status: Granted
- Kind: B2
- Application number: 17/892,549
- Filed: 2022-08-22
- Published: 2022-12-22
- Granted: 2024-02-13
- 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/US11898518B2

## 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 direct drive gas turbine engine mounted to an aircraft wing having a leading edge and a trailing edge, the direct drive 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, wherein the maximum outer flowpath diameter is defined at two or more rotating stages at a downstream-most end of the LP turbine, and wherein the gas turbine engine defines a fan to turbine diameter ratio of the maximum fan diameter to the maximum outer flowpath diameter less than or equal to 1.8, and wherein the fan to turbine diameter ratio over the fan pressure ratio is 0.90 or greater,
wherein a containment shield extends over the LP turbine from a first turbine rotor through a last turbine rotor along the longitudinal direction between the leading edge and the trailing edge of the aircraft wing,
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 a top dead center reference line along the circumferential direction, and
wherein the containment shield extends 60 degrees or less from the top dead center reference line along the circumferential direction.

### Claim 2 (depends on 1)

The gas turbine engine mounted to the aircraft wing of claim 1, wherein the fan pressure ratio defined by the fan assembly is between 1.0 and 1.8.

### Claim 3 (depends on 2)

The gas turbine engine mounted to the aircraft wing of claim 2, wherein the fan to turbine diameter ratio over the fan pressure ratio is 1.20 or greater at the fan pressure ratio of 1.50 or less.

### Claim 4 (depends on 2)

The gas turbine engine mounted to the aircraft wing of claim 2, wherein the fan to turbine diameter ratio over the fan pressure ratio is 1.05 or greater at the fan pressure ratio of between 1.50 to 1.60.

### Claim 5 (depends on 1)

The gas turbine engine mounted to the aircraft wing of claim 1, wherein the fan to turbine diameter ratio is 1.44 to 1.8.

### Claim 6 (depends on 1)

The gas turbine engine mounted to the aircraft wing 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 7 (depends on 1)

The gas turbine engine mounted to the aircraft wing of claim 1, wherein the LP turbine defines six or more stages of rotors.

### Claim 8 (depends on 1)

The gas turbine engine mounted to the aircraft wing of claim 1, wherein the LP turbine defines the maximum outer flowpath diameter at three or more rotating stages at the downstream-most end of the LP turbine.

### Claim 9 (depends on 1)

The gas turbine engine mounted to the aircraft wing of claim 1, 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 the downstream-most end of the LP turbine, and further wherein the LP turbine defines a turbine burst area inward of the aircraft wing along a latitudinal 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-most end, and at a second angle along a plane of rotation of the last turbine rotor toward the downstream-most end.

### Claim 10 (depends on 9)

The gas turbine engine mounted to the aircraft wing of claim 9, wherein the first angle of the turbine burst area is 15 degrees or less.

### Claim 11 (depends on 9)

The gas turbine engine mounted to the aircraft wing of claim 9, wherein the second angle of the turbine burst area is 15 degrees or less.

### Claim 12 (depends on 1)

The gas turbine engine mounted to the aircraft wing of claim 1, wherein the containment shield comprises:
a first containment shield extending at least along a transverse direction; and
a second containment shield extending at least along a circumferential direction.

### Claim 13 (depends on 12)

The gas turbine engine mounted to the aircraft wing of claim 12, wherein the second containment shield extends clockwise and counter-clockwise from the top dead center reference line.

### Claim 14 (depends on 12)

The gas turbine engine mounted to the aircraft wing of claim 12, wherein the first containment shield is spaced apart from the second containment shield.

### Claim 15 (depends on 1)

The gas turbine engine mounted to the aircraft wing of claim 1, wherein the containment shield comprises a ceramic matrix composite material.

### Claim 16 (depends on 1)

The gas turbine engine mounted to the aircraft wing of claim 1, wherein the containment shield extends clockwise and counter-clockwise from the top dead center reference line.

### Claim 17 (depends on 1)

The gas turbine engine mounted to the aircraft wing of claim 1, wherein the containment shield comprises a solid foamed synthetic polymer.