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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.

The necessity of avoiding the destructive and non-repeatable FSST (Full Scale Sled Test) makes it desirable to devise a cheaper and more repeatable method which can supplant this test procedure. This need developed the HCTD (HIC Component Testing Device) which is capable of providing conservative HIC results with higher repeatability. The computational model of the HCTD is validated against one of the tests conducted at CAMI with polyethylene foam. This validated model is used to conduct a series of tests with input parameters similar to the sled test to develop the correlation between the sled test and HCTD. This study hence concludes that a validated computational model of HCTD can be successfully utilized to address the HIC compliance issues for a foam padded surface.

An experimental study was performed to investigate large droplet dynamics in the vicinity of an airfoil. The investigation was conducted using the NASA Glenn Droplet Imaging Flow Tunnel (DrIFT). Mono-dispersed large droplets were released at the tunnel inlet and accelerated toward an airfoil that was mounted in the test section. The dynamic behavior of a droplet's encounter with the airfoil, which may involve droplet distortion, break-up, impingement and splashing, was recorded using a high-speed imaging system. The effects of the droplet size, tunnel velocity and airfoil configuration on the droplet dynamics were investigated in a parametric study. The droplet sizes used in the experimental study were 96 and 375 μm whereas tunnel velocities were varied from 80 to 130 mph. Three different airfoil geometries were used in the experimental study; a ‘clean’ and ‘iced’ airfoil, and a ‘clean’ three-element high-lift airfoil. The incidence angle of these airfoils was set to zero degrees.

A methodology for developing spray rigs for icing cloud simulation is presented. This methodology includes Computational Fluid Dynamics (CFD) analysis and icing tunnel experiments and was applied to design a spray rig system for a small airborne icing tanker. An in-house spray system was developed and tested in a laboratory to assess two commercially available nozzles - a single-jet type and a multi-jet type - which were capable of producing both FAR Part 25 Appendix C and SLD icing clouds. Spray rig characteristics evaluated during the laboratory tests included air and water flow rates as well as droplet size and distributions. The effects of airspeed and nozzle spacing on spray plume size and uniformity were investigated in a small icing tunnel facility with a two-nozzle spray rig. The experimental data were compared with three-dimensional numerical simulation results obtained with the FLUENT software.

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.

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.