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The work presented in this paper is part of a long-term research program to explore methods for improving bleed air system performance. Another objective of this research is to provide detailed experimental data for the development and validation of simulation tools used in the design and analysis of bleed air systems. A business jet wing was equipped with an inner-liner hot air ice protection system and was extensively instrumented for documenting system thermal performance. The wing was tested at the NASA Glenn Icing Research Tunnel (IRT) for representative in-flight icing conditions. Data obtained include bleed air supply and exhaust flow properties, wing leading edge skin temperatures, temperatures and pressures in the interior passages of the bleed air system, flow properties inside the piccolo tube, photos of run back ice shapes and ice shape traces. Selected experimental results for a warm hold icing condition are presented in this paper.

Ice particles shed from aircraft surfaces are a safety concern because they can damage aft-mounted engines and other aircraft components. Ice shedding is a random and complex phenomenon. The randomness of the ice fragment geometry, size, orientation and shed location in addition to potential particle breakup during flight poses considerable simulation challenges. Current ice shedding analysis tools have limited capabilities due to the lack of experimental aerodynamic coefficients for the forces and moments acting on the ice fragment. A methodology for simulating the shedding of large ice particles from aircraft surfaces was developed at Wichita State University. This methodology combines experimental aerodynamic characteristics of ice fragments, computational fluid dynamics, trajectory analysis and the Monte Carlo method to provide probability maps of shed particle footprints at desired locations.

This paper presents experimental methods for investigating large droplet impingement dynamics and for obtaining small and large water droplet impingement data. Droplet impingement visualization experiments conducted in the Goodrich Icing Wind Tunnel with a 21-in chord NACA 0012 airfoil demonstrated considerable droplet splashing during impingement. The tests were performed for speeds in the range 50 to 175 mph and with cloud median volumetric diameters in the range of 11 to 270 microns. Extensive large droplet impingement tests were conducted at the NASA Glenn Icing Research Tunnel (IRT). Impingement data were obtained for a range of airfoil sections including three 36-inch chord airfoils (MS(1)-0317, GLC-305, and NACA 652-415), a 57-inch chord Twin Otter horizontal tail section and 22.5-minute and 45-minute LEWICE glaze ice shapes for the Twin Otter tail section. Small droplet impingement tests were also conducted for selected test models.

Experimental studies were conducted to investigate the effect of tailplane icing on the aerodynamic characteristics of 15%-scale business jet aircraft. The simulated ice shapes selected for the experimental investigation included 9-min and 22.5-min smooth and rough LEWICE ice shapes and spoiler ice shapes. The height of the spoilers was sized to match the horns of the LEWICE shapes on the suction side of the horizontal tail. Tests were also conducted to investigate aerodynamic performance degradation due to ice roughness which was simulated with sandpaper. Six component force and moment measurements, elevator hinge moments, surface pressures, and boundary layer velocity profiles were obtained for a range of test conditions. Test conditions included AOA sweeps for Reynolds number in the range of 0.7 based on tail mean aerodynamic chord and elevator deflections in the range of -15 to +15 degrees.