See-Cheuk Wong

Icing Tunnel Experiments with a Hot Air Anti-Icing System

Experiments were conducted at the NASA Glenn Icing Research Tunnel with a 60-in chord business jet wing equipped with a hot air ice protection system. Tests were performed for warm hold, cold hold and descent icing conditions representative of in-flight icing. Bleed air flow and system parameters tested included range of hot air temperatures and mass flows, two piccolo configurations and insulation material. The main objective of the experimental study was to generate a detailed database for the development and validation of thermal and icing analysis codes. The wing model was instrumented with 205 temperature and pressure sensors to monitor wing leading edge skin temperatures and interior bleed air system flow properties. Selected experimental results for the warm hold case are presented in this paper to demonstrate the effects of piccolo hot air mass flow and temperature on system performance and runback ice accretions. The experimental results also include the effects of angle of attack, airspeed, piccolo hole pattern and leading edge skin insulation on wing skin temperature distributions for wet and dry external flow conditions.

Icing Tests of a Wing Model with a Hot-Air Ice Protection System

A business jet wing model equipped with a hot-air ice protection system was tested at the NASA Glenn Icing Research Tunnel for a range of external conditions representative of inflight icing. The 2D wing model was instrumented with 76 surface pressure taps and 205 temperature and pressure sensors to monitor wing leading edge skin temperatures and bleed-air system flow properties. The main objective of the investigation was to generate an experimental database for the development and validation of simulation tools used to design and analyze thermal ice protection system. Bleed-air system performance data and runback ice shapes were documented for two piccolo configurations and a range of bleed-air mass flows and temperatures for dry and wet external flow conditions. Selected experimental results for representative cold hold (V=110 kts, =3, Ts=-22F, MVD=29m, LWC=0.69 g/m 3 ) and descent (V=194 kts, =-1, Ts=-4F, MVD=25m, LWC=0.41 g/m 3 ) flight conditions are included in this paper. The experimental results presented demonstrate the variation in wing skin temperatures and other bleed air system parameters as a function of hot-air mass flow, hot-air temperature and piccolo configuration. The runback ice accretions documented showed that the ice front advanced upstream along the airfoil upper and lower surfaces as the thermal input to the bleed air system was reduced. Runback ice shape heights measured on the wing upper surface ranged from 0 to 0.26 inches.

Comparison of Experimental and Computational Ice Shapes for an Engine Inlet

range of -14F to 16F, and icing exposure times of 6, 10, 12, and 22.5 minutes. Ice shape traces were obtained at four inlet circumferential stations corresponding to 0 (inlet upper lip), 90, 180, and 270. Inlet leading edge ice shapes were computed with the LEWICE3D ice accretion code for the conditions investigated during the icing tunnel tests. The experimental and LEWICE3D ice shapes were compared to assess the performance of LEWICE3D in computing inlet ice accretion characteristics. Experimental and LEWICE3D ice shapes were found to be in good overall agreement, in terms of ice shape size, horn features, and icing limits, for six out of the nine cases presented in this paper. For the remaining three cases, however, the LEWICE3D shapes exhibited considerable differences with the experimental shapes.

Computation of Pre- and Post-Stall Flows over Single and Multi-Element Airfoils

The Navier-Stokes equations were used to predict the stall and post-stall behavior of five single element airfoils and a multi-element airfoil at low freestream Mach numbers. Results from a comprehensive numerical investigation are presented and compared with experimental force and pressure data. Two numerical schemes were used to integrate the governing equations and three turbulence models were applied to compute turbulent flow properties. The effect of grid resolution on the computational results is discussed. In most cases, the computed force and pressure coefficients obtained for a modified NACA 2412, a NACA 4412, a NLF 0414, a GA(W)-2, and an NREL S809 airfoil with 40% chord vented aileron and 40% chord spoiler flap, were in acceptable agreement with experimental data. These models, however, have shown limited success and cannot be easily extended to three-dimensi onal flows. Navier-Stokes based computer programs can be used to predict flows with large separation, such as near or even post-stall flow. The purpose of this study was to investigate the behavior of pre- and post-stall flow over single and multi-element airfoils by using two NavierStokes computer codes. Computational results presented include force coefficients, pressure and skin friction distributions and for selected cases Mach number contours which are used to identify regions of separated flow. In all cases the computational results are compared with available experimental data.

Computational and Experimental Investigation of a Spray Rig for Airborne Icing Tankers

Computational and experimental investigations were conducted to evaluate the performance of a spray rig for airborne icing tanker applications. Tests were performed at the Wichita State University (WSU) 7-ft x 10-ft wind tunnel facility to assess the effects of spray nozzle number density, pattern and spacing on spray cloud plume size and uniformity. Air-atomizing multi-jet nozzles and vortex generators were incorporated into the spray rig design to enhance cloud size and uniformity. Computational fluid dynamics (CFD) analyses were conducted prior to the wind tunnel tests to evaluate the effects of vortex generators and nozzle support structure on spray cloud droplet dispersal. FAR Part 25 Appendix C and SLD spray cloud conditions were generated with a computer controlled spray system developed at WSU. Spray plume sizes were documented at an airspeed of 155 mph using a laser sheet imaging (LSI) technique. Spray cloud liquid water content (LWC) was measured with a CSIRO-King probe installed at the diffuser section of the tunnel. Three nozzle arrangements with 7, 11 and 12 nozzles placed within a circle of 24-in radius were investigated. Spray plume sizes measured at a location 10.5-ft downstream of the nozzles were approximately 5-ft in diameter for all three nozzle arrangements tested. The 11-nozzle spray rig tested provided the best overall cloud uniformity. Spray cloud LWC for the 11nozzle spray rig was 0.13 g/m 3 for a cloud median volumetric diameter (MVD) of 40 microns, and 0.47 g/m 3 for a cloud MVD of 50 microns. For large droplet spray clouds, the LWC varied from 1.20 to 1.50 g/m 3 as the cloud MVD was increased from 100 to 160 microns.