This paper discusses the application of pressure sensitive paint using laser-based excitation for measurement of the upper surface pressure distribution on the tips of rotor blades in hover and simulated forward flight. The testing was conducted in the Rotor Test Cell and the 14-by 22-ft Subsonic Tunnel at the NASA Langley Research Center on the General Rotor Model System (GRMS) test stand. The Mach-scaled rotor contained three chordwise rows of dynamic pressure transducers for comparison with PSP measurements. The rotor had an 11 ft 1 in. diameter, 5.45 in. main chord and a swept, tapered tip. Three thrust conditions were examined in hover, C T = 0.004, 0.006 and 0.008. In forward flight, an additional thrust condition, C T = 0.010 was also examined. All four thrust conditions in forward flight were conducted at an advance ratio of 0.35.

This effort examined the potential of multi-winglets for the reduction of induced drag without increasing the span of aircraft wings. Wind tunnel models were constructed using a NACA 0012 airfoil section for the untwisted, rectangular wing and flat plates for the winglets. Testing of the configurations occurred over a range of Reynolds numbers from 161,000 to 300,000. Wind tunnel balances provided lift and drag measurements, and laser flow visualization obtained wingtip vortex information.The Cobalt 60 unstructured solver generated flow simulations of the experimental configuration via solution of the Euler equations of motion. The results show that certain multi-winglet configurations reduced the wing induced drag and improved L/D by 15-30% compared with the baseline 0012 wing. A substantial increase in lift curve slope occurs with dihedral spread of winglets set at zero incidence relative to the wing. Dihedral spread also distributes the tip vortex. These observations supplement previous results on drag reduction due to lift reorientation with twisted winglets set at negative incidence.

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This paper will present details of a Pressure Sensitive Paint (PSP) system for measuring global surface pressures on the tips of rotorcraft blades in simulated forward flight at the 14-by 22-Foot Subsonic Tunnel at the NASA Langley Research Center. The system was designed to use a pulsed laser as an excitation source and PSP data was collected using the lifetime-based approach. With the higher intensity of the laser, this allowed PSP images to be acquired using a single laser pulse, resulting in the collection of crisp images that can be used to determine blade pressure at a specific instant in time. This is extremely important in rotorcraft applications as the blades experience dramatically different flow fields depending on their position. Testing of the system was performed using the U.S. Army General Rotor Model System equipped with four identical blades. Two of the blades were instrumented with pressure transducers to allow for comparison of the results obtained from the PSP. Preliminary results show that the PSP agrees both qualitatively and quantitatively with both the expected results as well as with the pressure transducers. Several areas of improvement have been indentified and are currently being developed for future testing.