Marianne Monastero

Validation of 3-D Ice Accretion Measurement Methodology for Experimental Aerodynamic Simulation

Determining the adverse aerodynamic effects due to ice accretion often relies on dry-air wind-tunnel testing of artificial, or simulated, ice shapes. Recent developments in ice-accretion documentation methods have yielded a laser-scanning capability that can measure highly three-dimensional (3-D) features of ice accreted in icing wind tunnels. The objective of this paper was to evaluate the aerodynamic accuracy of ice-accretion simulations generated from laser-scan data. Ice-accretion tests were conducted in the NASA Icing Research Tunnel using an 18-in. chord, two-dimensional (2-D) straight wing with NACA 23012 airfoil section. For six ice-accretion cases, a 3-D laser scan was performed to document the ice geometry prior to the molding process. Aerodynamic performance testing was conducted at the University of Illinois low-speed wind tunnel at a Reynolds number of 1.8 10(exp 6) and a Mach number of 0.18 with an 18-in. chord NACA 23012 airfoil model that was designed to accommodate the artificial ice shapes. The ice-accretion molds were used to fabricate one set of artificial ice shapes from polyurethane castings. The laser-scan data were used to fabricate another set of artificial ice shapes using rapid prototype manufacturing such as stereolithography. The iced-airfoil results with both sets of artificial ice shapes were compared to evaluate the aerodynamic simulation accuracy of the laser-scan data. For five of the six ice-accretion cases, there was excellent agreement in the iced-airfoil aerodynamic performance between the casting and laser-scan based simulations. For example, typical differences in iced-airfoil maximum lift coefficient were less than 3 percent with corresponding differences in stall angle of approximately 1 deg or less. The aerodynamic simulation accuracy reported in this paper has demonstrated the combined accuracy of the laser-scan and rapid-prototype manufacturing approach to simulating ice accretion for a NACA 23012 airfoil. For several of the ice-accretion cases tested, the aerodynamics is known to depend upon the small, three-dimensional features of the ice. These data show that the laser-scan and rapid-prototype manufacturing approach is capable of replicating these ice features within the reported accuracies of the laser-scan measurement and rapid-prototyping method; thus providing a new capability for high-fidelity ice-accretion documentation and artificial ice-shape fabrication for icing research.

Aerodynamics of a Swept Wing with Ice Accretion at Low Reynolds Number

This paper discusses the use of several experimental techniques to investigate the performance and flowfield of a swept wing with leading-edge ice at low Reynolds numbers. Force balance measurements were made over a range of angles of attack and Reynolds numbers. Surface oil visualization and pressure sensitive paint were used to investigate the flow over the surface of the wing, while five-hole probe wake surveys were used to investigate the wake. The flowfield of the iced swept wing was dominated by a leading-edge vortex that formed at low angles of attack due to separation from the tip of the ice shape, while for the clean wing a vortex did not form until higher angles of attack. The effect of Reynolds number on the performance and flowfield was also investigated.

Experimental Study of Splitter Plates for Use with Semispan Wing Models

A common technique when testing a three-dimensional wing in a wind tunnel is to use a semispan, also known as a half-span, wing as opposed to a complete aircraft model. The semispan technique has several important advantages over a model of a complete aircraft. Semispan models can typically achieve double the Reynolds number of the similar full-span model due to the increased size. Additional benefits of the large model size include improved model strength, higher geometric fidelity and more room for instrumentation. The semispan model can also typically be constructed for less cost if complex features such as high-lift systems or flow through nacelles are required. 1 The primary drawbacks to the semispan model include increased wind tunnel wall effects due to the large model size, and more importantly, the aerodynamic influence of the specific mounting method. Unlike a full-span model, a semispan wing does not require any sting or support strut, and there are several ways in which a semispan wing is typically mounted in a wind tunnel. The most straightforward method is to simply mount the wing directly to the tunnel wall. A semispan model is typically one side of a symmetric geometry so ideally the tunnel wall would act as a reflection plane boundary condition. Unfortunately, there are several ways in which the tunnel floor boundary layer can influence the aerodynamics of the semispan model and invalidate the reflection plane boundary condition. First, the loading near the root of the semispan wing is reduced due to the low dynamic pressure in the tunnel floor boundary layer. Secondly, the tunnel floor boundary layer can separate near the wing leading edge resulting in the formation of a horseshoe vortex system. Finally, the adverse pressure gradient along the surface of the wing can cause the tunnel floor boundary layer to separate anywhere along the aft section of the wing root. 1,2,3,4 These modifications to the flowfield resulting from the tunnel floor boundary layer can lead to a reduction in the lift curve slope, increased zero lift incidence and increased drag on the semispan wing compared to the full-span. 2,3,5

Validation of 3-D Ice Accretion Measurement Methodology Using Pressure-Sensitive Paint

Accurate representation of ice accretions is important to the study and understanding of aircraft icing. For research and certification purposes, replicas of ice accretions generated from an icing wind tunnel are created to perform aerodynamic tests in dry-air wind tunnels and in flight. Three-dimensional ice-shape features necessitate aerodynamic observations over the entire surface of the wing or airfoil to fully understand the flow behavior. The pressure-sensitive paint (PSP) technique allows pressure coefficient (Cp) data to be obtained over a larger area and with a greater resolution than is possible by solely using the pressure tap method. Three-dimensional Cp measurements obtained as the result of this work both further the understanding of iced-airfoil aerodynamics and provide a validation for a recently developed method of recording ice accretions. Results show the ability of the PSP technique implemented in this experimental study to resolve aerodynamic differences between ice shapes made from the new 3-D ice accretion measurement methodology recently developed by the NASA Icing Research Branch and the current mold and casting method. PSP tests were performed to further investigate the aerodynamics of ice shapes that showed performance variations between the replicas made from the new and current methods. In general, the PSP results agreed with the two-dimensional data obtained from pressure taps while providing additional pressure information along the span. Aerodynamic and PSP agreement between the shapes generated from the two methods was relatively good, but not to the extent expected. Results show the general behaviors agree, with discrepancies in the quantitative data. These discrepancies are shown to be the result of geometric differences in the ice shapes made from the two methods. The 3-D ice accretion measurement methodology was validated for the two tested shapes and possible improvements to the PSP implementation are suggested.

Three-Dimensional Ice-Accretion Measurement Methodology for Experimental Aerodynamic Simulation

Determining the adverse aerodynamic effects due to ice accretion often relies on dry-air wind-tunnel testing of artificial, or simulated, ice shapes. Recent developments in ice-accretion documentat...