Gas turbine engine
US10598022B1 · kind B1 · utility
Assignee
Inventors
Key dates
| Filing date | Jun 25, 2019 |
| Grant date | Mar 24, 2020 |
| Priority date | — |
| Expiry date | Jun 25, 2039 |
Classification
- Technology area (CPC Y)Emerging Cross-Sectional Technologies
- CPC primaryY02T50/60
- WIPO fieldEngines, pumps, turbines
- WIPO sectorMechanical engineering
Abstract
A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising: a compressor system comprising a first, lower pressure, compressor (14), and a second, higher pressure, compressor (15); and an outer core casing (70) surrounding the compressor system. The gas turbine engine further comprises a fan (23) located upstream of the engine core (11), the fan comprising a plurality of fan blades. The outer core casing comprises: a first flange connection (60) arranged to allow separation of the outer core casing (70) at an axial position of the first flange connection (60), the first flange connection (60) having a first flange radius (104), wherein the first flange connection (60) is the first flange connection that is downstream of an axial position defined by the axial midpoint between the mid-span axial location on the trailing edge of the most downstream aerofoil of the first compressor (14) and the mid-span axial location on the leading edge of the most upstream aerofoil of the second compressor (15). A fan blade mass ratio of: is equal to or less than 19.0 mm/lb.
Source: USPTO / EPO open patent data. Objective bibliographic and citation counts.