Gas turbine engine system providing simulated boundary layer thickness increase
US8209953B2 · kind B2 · utility
12Cited by
111References
8Claims
0Family size
Assignee
Inventors
Key dates
| Filing date | Dec 9, 2010 |
| Grant date | Jul 3, 2012 |
| Priority date | — |
| Expiry date | Dec 9, 2030 |
Classification
- Technology area (CPC Y)Emerging Cross-Sectional Technologies
- CPC primaryY10T137/0536
- WIPO fieldEngines, pumps, turbines
- WIPO sectorMechanical engineering
Abstract
A gas turbine engine system for an aircraft includes a nacelle having a fan cowl with an inlet lip section and a core cowl, at least one compressor and at least one turbine, at least one combustor between the compressor and the turbine, a bleed passage, and a controller. The bleed passage includes an inlet for receiving a bleed airflow and an outlet that discharges the bleed airflow in an upstream direction from the outlet. The controller identifies an operability condition and selectively introduces the bleed airflow near a boundary layer of the inlet lip section in response to the operability condition.
Source: USPTO / EPO open patent data. Objective bibliographic and citation counts.