Hsiung-Wei Yeong

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.

Experimental and Computational Investigation of Ice Shedding from Aircraft Surfaces

Ice particles shed from aircraft surfaces are a safety concern because they can damage aft mounted engines and other aircraft components. Current particle trajectory simulation methods have limited capabilities in predicting ice fragment trajectories. This is due to the random characteristics of the shed particles and lack of experimental aerodynamic data for ice fragments. This paper describes a new effort in developing a methodology for computing the trajectories of large ice particles. The 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. The methodology was applied to compute the trajectories of a square ice fragment and a rectangular plate shed in a three-dimensional uniform flowfield. Monte Carlo simulations were also performed where ice fragment parameters such as length, thickness, lift and drag coefficients, initial orientation, etc. were randomly changed and probability maps were computed.

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.