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Results from a two-dimensional computer model developed at Wichita State University (WSU) for bleed air system analysis are compared with experimental data from icing tunnel tests performed with a wing model equipped with a hot air ice protection system. The computer model combines a commercial Navier-Stokes flow solver with a steady-state thermodynamic analysis model that applies internal flow heat transfer correlations to compute wing leading edge skin temperatures and the location and extent of the runback ice. The icing tunnel data used in the validation of the computer model were obtained at the NASA Icing Research Tunnel using representative in-flight icing conditions and a range of bleed air system mass flows and hot air temperatures. Correlation between experiment and analysis was good for most of the test cases used to assess the performance of the simulation model.

An experimental methodology for investigating the effects of anti-icing fluids is presented in this paper. A wing model was designed, fabricated, and instrumented for testing anti-icing fluids in a wind tunnel facility. In addition, a video capturing method was developed and used to document fluid behavior during simulated takeoff tests. The experiments were performed at the Wichita State University 2.13-m by 3.05-m (7-ft by 10-ft) wind tunnel facility with two pseudoplastic fluids representative of Type IV anti-icing fluids. The experimental data obtained included fluid wave propagation speeds, chordwise fluid thickness distributions as a function of time, and boundary layer velocity profiles for the clean and fluid contaminated wing model at select chordwise stations. During simulated takeoffs with initial fluid depths of either 4 mm or 2 mm, the fluids were observed to thin in the forward (upstream) regions of the wing model and accumulate in the aft regions.

Two-dimensional and three-dimensional leading edge ice shapes for a finite wing model computed with the NASA Glenn LEWICE 2.0 and LEWICE3D Version 2 ice accretion codes are compared with experimental data from icing tunnel tests. The wing model had 28° leading edge sweep angle, 1.52-m (60-in) semispan and an airfoil section representative of business jet wings. Experimental wing leading edge ice shapes were obtained at the NASA Glenn Icing Research Tunnel (IRT) for six icing conditions. Tests conditions included angles of attack of 4° and 6°, airspeeds ranging from 67.06 m/s (150 mph) to 111.76 m/s (250 mph), static air temperatures in the range of -11.28°C (11.7°F) to -2.78°C (27°F), liquid water contents of 0.46 g/m₃, 0.51 g/m₃, and 0.68 g/m₃, and median volumetric diameters of 14.5 μm and 20 μm.

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.

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.

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.

An experiment was conducted in the Icing Research Tunnel (IRT) at NASA Glenn Research Center to obtain castings of ice accretions formed on a 28° swept GLC-305 airfoil that is representative of a modern business aircraft wing. Because of the complexity of the casting process, the airfoil was designed with three removable leading edges covering the whole span. Ice accretions were obtained at six icing conditions. After the ice was accreted, the leading edges were detached from the airfoil and moved to a cold room. Molds of the ice accretions were obtained, and from them, urethane castings were fabricated. This experiment is the icing test of a two-part experiment to study the aerodynamic effects of ice accretions.

Collection efficiency and ice accretion calculations have been made for a series of business jet horizontal tail configurations using a three-dimensional panel code, an adaptive grid code, and the NASA Glenn LEWICE3D grid based ice accretion code. The horizontal tail models included two full scale wing tips and a 25% scale model. Flow solutions for the horizontal tails were generated using the PMARC panel code. Grids used in the ice accretion calculations were generated using the adaptive grid code ICEGRID. The LEWICE3D grid based ice accretion program was used to calculate impingement efficiency and ice shapes. Ice shapes typifying rime and mixed icing conditions were generated for a 30 minute hold condition. All calculations were performed on an SGI Octane computer. The results have been compared to experimental flow and impingement data.

An experimental study was conducted to investigate ice-adhesion on clean and coated aluminum surfaces. A test apparatus using the parallel plate linear shear technique was designed along with a data acquisition system for conducting the tests and recording the experimental data. A low pulling rate was applied to specially prepared test specimens for measuring the strength of ice adhesion for a range of test conditions. The effects of surface roughness, surface contamination, and water impurity on ice adhesion were investigated. In addition, tests were conducted to evaluate the effectiveness of a low ice-adhesion coating applied to aluminum test specimens. The results obtained showed that the bond between ice and metal was considerably lower for tap water than for distilled water. For the clean and coated aluminum surfaces the strength of ice adhesion varied with specimen roughness. However, no clear trend was established between ice adhesion strength and surface roughness.

An experimental investigation on the effects of Gurney flaps on a reflection plane model was conducted. Two sizes of Gurney flaps were tested on a series of configurations which included a tapered wing (with a NLF-0215 airfoil section), a fuselage, a nacelle, and their permutations. The tests were conducted at a Reynolds number of 1.0 million based on mean chord. Results indicated that lift and drag were increased upon using the Gurney flaps; lift to drag and lift squared to drag ratios were also increased. In particular, the lift to drag ratio for the complete “airplane” was almost the same with or without a small Gurney flap. Pitching moment became more negative (nose-down) with the Gurney flap, and positive (nose-up) with the addition of the fuselage.

The effect of Gurney flaps on three-dimensional wings was investigated in the 7x10 feet low speed wind tunnel. There have been a number of studies on Gurney flaps in recent years. However, these studies have been limited to two-dimensional airfoils. A comprehensive investigation on the effect of Gurney flaps for a wide range of configurations and test conditions was conducted at Wichita State University. In this part of the investigation, straight and tapered three-dimensional wings with Natural Laminar Flow (NLF) airfoil sections were tested. Gurney flaps spanning 4.5, 3.0, and 1.5 feet were tested on a straight NLF wing of 5 feet span. Compared to the clean wing, the 4.5 feet span 0.017c and 0.033c height Gurney flaps increased the maximum lift coefficient by 17% and 22%, respectively. The increase in maximum lift coefficient was proportionately smaller with the shorter span Gurney flaps.

The effect of Gurney flaps on a NACA 0011 airfoil was investigated. Gurney flaps provide a substantial increase in lift while the penalty in drag is small. With the Gurney flap, the airfoil pressure distribution shows increased suction on the upper surface and higher pressure on the lower surface compared to the clean airfoil. This change in pressure is most profound on the lower surface just in front of the Gurney flap. Since separation occurs on the upper surface upon stall, this higher pressure condition on the lower surface continues into the post-stall regime. Thus, the NACA 0011 airfoil with Gurney flaps generates lift coefficients greater than one even under post-stall conditions.

The flow about multi-element airfoil configurations is investigated using the unsteady Reynolds averaged Navier-Stokes equations. An explicit scheme is used to advance the solution in time while a finite difference scheme is applied to discretize the flux terms. An algebraic and two one-equation turbulence models are used to model turbulence. The domain about each multi-element airfoil is discretized with structured Chimera grids. The multi-element configurations presented in this paper include two airfoils with slotted flaps and an airfoil with a 50% chord vented aileron deflected at 90 degrees. Subsonic flow computations are performed for attached and separated flow conditions. The computational results obtained with the CRTVD code developed at Wichita State University are in good correlation with wind tunnel data and with computational results obtained with the INS2D computer code developed at NASA Ames research center.

Flow field computations and, in particular, that of pressure, skin friction, and heat transfer (for high speed flights) are the primary parameters in the design of aerospace vehicles. Most computational schemes based on either the inviscid Euler equations or various forms of the Navier-Stokes equations are remarkably accurate in the predictions of pressure distributions. However, computations of skin friction and heat transfer particularly at high speeds have been a source of considerable difficulty. Problems arise not only due to the grid resolution but also due to the particular numerical scheme employed. To address the difficulty associated with accurate computations of the velocity and temperature gradients, a comparative investigation of several Navier- Stokes codes is undertaken. Previous studies with regard to the effect of grid resolution are incorporated into the current investigation.

A Non-Planar Vortex Lattice Method (VLM) has been combined with a Two-Dimensional Surface Panel Method for computing the aerodynamic characteristics of finite wings in incompressible inviscid flow. This numerical model can be applied to wings with thickness and arbitrary planform, and requires very little computing time when compared to Three-Dimensional Surface Panel Methods currently in use. The formulation of the present method is described in detail, and results from its application to three wing configurations are presented. The results obtained using the present method are compared with results obtained using the VSAERO code and with experimental data. Good correlation is demonstrated in all cases.