Dynamic Stall of Swept and Unswept Oscillating Wings,
Abstract
Recent unsteady tests on oscillating tunnel-spanning wings (representative of full scale helicopter blades) have shown that the dynamic stall phenomenon for a swept wing model is significantly different from that for an unswept wing. Several critical measurements and calculations relative to the behavior of the surface flow were made, including chordwise wave speed of the stalling vortex, the degree of pitch rate dependence of the vortex inception angle, and the ability of the cosine law for sweep to normalize unsteady post-stall behavior, even after reattachment occurs. The vortex chordwise wave speed was found to be insensitive to sweep angle over the forward blade region, but was approximately half as large over the aft region for a sweep angle of 30 deg, relative to the unswept wing. It was also found that the vortex inception angle (which can be associated with dynamic stall inception) was virtually unaffected by changes in sweep angle. However, a change in Mach number (measured normal to the leading edge) from 0.3 to 0.4 caused a drop in vortex inception angle of about 4 deg, comparable to the change in quasi-steady stall angle for the same Mach number range. It was also found that the vortex inception angle varies linearly with reduced frequency, and appears to be independent of pitch rate at constant reduced frequency. The traditional cosine law normalization for sweep is not applicable above the stall angle, either for steady or for unsteady behavior. As a corollary to this, it was determined that sweep causes a dynamic displacement of the lift response such that the unstalled portion of a lift force coefficient, airfoil angle hysteresis loop (both prestall and post-stall) for the swept wing lies below that for the unswept wing.
Document Details
- Document Type
- Technical Report
- Publication Date
- Nov 01, 1985
- Accession Number
- ADP005010
Entities
People
- Franklin O. Carta
Organizations
- United Technologies Corporation