Gas turbine engine fan
US-2017138306-A1 · May 18, 2017 · US
Large-scale bypass fan configuration for turbine engine core and bypass flows
US11339713B2 · US · B2
| Field | Value |
|---|---|
| Publication number | US-11339713-B2 |
| Application number | US-201916398742-A |
| Country | US |
| Kind code | B2 |
| Filing date | Apr 30, 2019 |
| Priority date | Dec 21, 2018 |
| Publication date | May 24, 2022 |
| Grant date | May 24, 2022 |
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- Title
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Abstract
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A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, a core shaft, and a core exhaust nozzle, the core exhaust nozzle having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit; a fan including a plurality of fan blades; and a nacelle surrounding the fan and the engine core and defining a bypass duct, the bypass duct including a bypass exhaust nozzle, the bypass exhaust nozzle having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit;wherein a bypass to core ratio of:bypassexhaustnozzlepressureratiocoreexhaustnozzlepressureratiois configured to be in the range from 1.1 to 2.0 under aircraft cruise conditions.
First claim
Opening claim text (preview).
The invention claimed is: 1. A gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, a core shaft connecting the turbine to the compressor, and a core exhaust nozzle having a core exhaust nozzle exit, the core exhaust nozzle having a core exhaust nozzle pressure ratio calculated using total pressure at the core exhaust nozzle exit; the turbine comprises a lowest pressure turbine stage having a row of rotor blades, each of the rotor blades extending radially and having a leading edge and a trailing edge; a fan located upstream of the engine core, the fan comprising a plurality of fan blades, wherein a fan tip radius of the fan is measured between a centreline of the gas turbine engine and an outermost tip of one of the plurality of fan blades at its leading edge; and a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside of the engine core, the bypass duct comprising a bypass exhaust nozzle having a bypass exhaust nozzle exit, the bypass exhaust nozzle having an outer radius measured as a radial distance between the centreline of the gas turbine engine and an inner surface of the nacelle at an axial position of a rearmost tip of the nacelle, wherein an outer bypass to fan ratio of: the outer radius of the bypass exhaust nozzle the fan tip radius is in the range from 0.95 to 1.00, the fan tip radius being in the range from 110 cm to 150 cm, and the bypass exhaust nozzle having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass exhaust nozzle exit; wherein a bypass to core ratio of: bypass exhaust nozzle pressure ratio core exhaust nozzle pressure ratio is configured to be in the range from 1.1 to 1.4 under aircraft cruise conditions, and wherein a fan-turbine radius difference is defined as a radial distance between: a point on a circle swept by a radially outer tip of the trailing edge of one of the rotor blades of the lowest pressure turbine stage; and a point on a circle swept by the outermost tip of the leading edge of the one of the plurality of fan blades and a fan speed to fan-turbine radius ratio defined as: a maximum take off rotational speed of the fan ( in rpm ) fan turbine radius difference ( in mm ) is in the range between 2.9 rpm/mm to 3.8 rpm/mm. 2. The gas turbine engine of claim 1 , wherein the total pressure at the bypass exhaust nozzle exit is determined at an exit plane of the bypass exhaust nozzle, the exit plane extending from the rearmost tip of the nacelle towards the centreline of the gas turbine engine. 3. The gas turbine engine of claim 1 , wherein the engine core comprises a casing, and wherein the total pressure at the core exhaust nozzle exit is determined at an
Assignees
Inventors
Classifications
for aircraft propulsion, e.g. jet engines · CPC title
characterised by construction · CPC title
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) · CPC title
- F02K3/06Primary
with front fan · CPC title
Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto (rocket nozzles F02K9/97) · CPC title
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- What does patent US11339713B2 cover?
- A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, a core shaft, and a core exhaust nozzle, the core exhaust nozzle having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit; a fan including a plurality of fan blades; and a nacelle surrounding the fan and the engine core and defining a bypass duct, the bypass…
- Who is the assignee on this patent?
- Rolls Royce Plc
- What technology area does this patent fall under?
- Primary CPC classification F02K3/06. Mapped technology areas include Mechanical Engineering.
- When was this patent published?
- Publication date Tue May 24 2022 00:00:00 GMT+0000 (Coordinated Universal Time) (B2). Legal status and post-grant events are not shown on this page.
- What related patents are in patentsdb?
- We list 12 related publications on this page (citations in our corpus or others sharing the same primary CPC).