Wing Pressure Distribution and Boundary Layer Data Obtained from C-5A Flight Testing.
Abstract
Analyses have been conducted of wing pressure distribution and boundary layer data obtained from flight tests of a C-5 airplane at Reynolds numbers from 35 to 90 million. Results showed that shock locations at high subsonic Mach numbers are as much as 10 to 12 percent chord aft of those measured in previous wind tunnel tests at 7.4 million Reynolds number. No consistent variation in shock location with Reynolds number within the range covered by the flight data can be detected, however. Considered of the boundary layer data in conjunction with the pressure measurements would indicate that this absence of scale effects at high Reynolds number results from the fact that trailing-edge separation is suppressed to the extent that separation at the shock is the dominant factor leading to flow breakdown.
Document Details
- Document Type
- Technical Report
- Publication Date
- Oct 01, 1974
- Accession Number
- ADA005324
Entities
People
- Jones F. Cahill
- William A. Stevens