Thrust force generation device and aircraft
US-10752369-B2 · Aug 25, 2020 · US
Aero engine flow rate
US10882633B2 · US · B2
| Field | Value |
|---|---|
| Publication number | US-10882633-B2 |
| Application number | US-202016741005-A |
| Country | US |
| Kind code | B2 |
| Filing date | Jan 13, 2020 |
| Priority date | Dec 21, 2018 |
| Publication date | Jan 5, 2021 |
| Grant date | Jan 5, 2021 |
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- Title
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Abstract
Official abstract text for this publication.
A gas turbine engine of an aircraft includes: an engine core having a turbine including a lowest pressure rotor stage, a turbine diameter, a fan including a plurality of fan blades extending from a hub, an annular fan face at a leading edge of the fan; wherein a downstream blockage ratio of: the turbine diameter at an axial location of the lowest pressure rotor stage a distance from a ground plane to the wing is in the range from 0.2 to 0.3.
First claim
Opening claim text (preview).
The invention claimed is: 1. A gas turbine engine for an aircraft mounted beneath a wing of the aircraft, the engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine comprising a lowest pressure rotor stage, the turbine having a turbine diameter; a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein a downstream blockage ratio of: the turbine diameter at an axial location of the lowest pressure rotor stage a distance from a ground plane to the wing is in a range from 0.2 to 0.3. 2. The gas turbine engine of claim 1 , wherein the downstream blockage ratio is in the range from 0.20 to 0.29. 3. The gas turbine engine of claim 1 , wherein the downstream blockage ratio is in the range from 0.22 to 0.28. 4. The gas turbine engine of claim 1 , wherein the distance between the ground plane and the wing is measured from the ground plane to a center point of a leading edge of the wing. 5. The gas turbine engine of claim 1 , wherein the distance between the ground plane and the wing is measured along a line perpendicular to the ground plane and passing through and perpendicular to an axial centreline of the engine. 6. The gas turbine engine of any claim 1 , wherein the turbine diameter at the axial location of the lowest pressure rotor stage is measured adjacent blade tips of rotor blades of the lowest pressure rotor stage. 7. The gas turbine engine of claim 1 , wherein the turbine diameter at the axial location of the lowest pressure rotor stage is in a range from 70 cm to 170 cm, and optionally wherein: (i) a fan tip radius of the fan in a range from 110 cm to 150 cm, and the turbine diameter at the axial location of the lowest pressure rotor stage is in a range from 100 cm to 120 cm; or (ii) a fan tip radius of the fan is in a range from 155 cm to 200 cm, and the turbine diameter at the axial location of the lowest pressure rotor stage is in a range from 120 cm to 170 cm. 8. The gas turbine engine of any claim 1 , wherein the hub and the plurality of the fan blades of the fan together define a fan face having a fan tip radius, and the engine has an engine length, and wherein: an engine ratio of: ( 2 × the fan tip radius / the engine length ) the downstream blockage ratio is in a range from 2.5 to 4. 9. The gas turbine engine of claim 8 wherein the engine ratio is in a range from 2.5 to 4.0, optionally 2.7 to 3.7. 10. The gas turbine engine of claim 8 wherein the engine ratio is greater than 2.5. 11. The gas turbine engine of claim 8 , wherein the engine ratio is greater than 3.0. 12. The gas turbine engine of claim 8 , wherein the engine length is measured as an axial distance between a forward region of the fan and a rearward region of the turbine. 13. The gas turbine engine of claim 8 , wherein the engine length is measured along a centerline of the engine from an axial position of an intersection of a leading edge of the fan and the hub of the fan to a trailing edge mean radius point of one of rotor blades provided
Assignees
Inventors
Classifications
- F02K3/06Primary
with front fan · CPC title
within, or attached to, wings · CPC title
Mathematical features · CPC title
specially adapted for the fan of turbofan engines · CPC title
Rotors for more than one axial stage, e.g. of drum or multiple disc type; Details thereof, e.g. shafts, shaft connections {(F01D5/022, F01D5/023 take precedence)} · CPC title
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- What does patent US10882633B2 cover?
- A gas turbine engine of an aircraft includes: an engine core having a turbine including a lowest pressure rotor stage, a turbine diameter, a fan including a plurality of fan blades extending from a hub, an annular fan face at a leading edge of the fan; wherein a downstream blockage ratio of: the turbine diameter at …
- 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 Jan 05 2021 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 3 related publications on this page (citations in our corpus or others sharing the same primary CPC).