Development of Flow Distortions in a Full-Scale Nacelle Inlet Mach Numbers 0.63 and 1.6 to 2.0
Abstract
A typical full-scale nose inlet was tested at Mach numbers 0.63 and 1.6 to 2.0 to determine the nature of flow-distortion development throughout the subsonic diffuser. Inlet design variables studied included 14 deg and 17 deg internal cow-lip angles, conical compression surfaces with and without boundary-layer removal slots, and cone tip translation. Angles of attack to -8 deg were investigate at Mach 2.0. At zero angle of attack, changes in cowl angle or compression-surface bleed had little effect on the distortion at the diffuser exit, his value being fairly well predicted by calculations of the turbulent-flow profile. Operation at other angles of attack, however, consistently pro used diffuser-exit distortions larger than the calculated values.
Document Details
- Document Type
- Technical Report
- Publication Date
- Oct 22, 1956
- Accession Number
- AD0111918
Entities
People
- Bruce G. Chiccine
- Thomas G. Piercy
Organizations
- National Aeronautics and Space Administration