Search Results

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.

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 present scenario of increasing the production rates in drilling demands high speed drilling while at the same time maintaining the quality of the drilled holes. Of the many factors that affect the high-speed drilling process, such as speed, feed rate, material of drill and work piece, and drill geometry this study attempts only to study the effect of drill geometry in high-speed drilling of aluminum sheet metal. In the Experiments conducted, different speeds, feed rates and drill bits of varying geometry are utilized in order to study their effects on hole quality as it relates to hole diameter and burr formation. Also the variation of thrust and torque with increase in speed over a speed range of 6,000 to 30,000 rpm has been studied.

An investigation was performed to evaluate a novel and inexpensive wind tunnel model mount for dynamic aerodynamic testing. A computer analysis code was developed to identify the dimensions of the control surface needed to produce a desired pitching motion for a delta wing. The code was then used to design and build a dynamic model apparatus that was evaluated in a low speed wind tunnel at Wichita State University. The dynamic model mount and control were evaluated for a variety of motions, including constant pitch rate ramps, constant frequency oscillations and impulse or step inputs. Results from the ramp and oscillation test indicated the system is very responsive and capable of a wide range of motion.

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.

A new class of aerodynamic control devices have recently been designed specifically for wind turbine applications. These new controls were tested to evaluate their effectiveness in modulating wind turbine power output and for slowing or stopping a wind turbine in high wind or loss of generator situations. While these control devices were developed specifically for wind turbine applications, there exists the possibility that alternate aviation uses exist. In particular, these trailing-edge control devices were evaluated for reducing aircraft landing distances, generating rapid rates of descent, deep stall or spin recovery and for high angle of attack control.

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 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.

A basic one-component Doppler Global Velocimetry (DGV) system has been developed at Wichita State University. This system was evaluated on a round axisymmetric jet. The results are compared with measurements made using traditional Constant Temperature Anemometry (CTA) and Preston tube measurements at 3.2, 9.6 and 16.0 jet exit diameters downstream and along the jet centerline. The DGV results show similar trends to these measurements. Software corrections for camera misalignment, optical distortions, and laser frequency variations were necessary to assure data quality. Results indicate good agreement between the DGV and CTA measurements exist.

This paper compares the environmental durability of currently used as well as some of the potential aircraft aluminum alloy skin materials. A simple test was developed to evaluate the environmental durability by simultaneous application of fatigue loads and aerated salt water attack. Alclad alloys showed excellent resistance to corrosion fatigue. On the other hand, 7475 alloy and new potential material 6013 alloy experienced inter-granular corrosion at the fatigue crack origin area. The test results also indicated the significance of corrosion preventive coatings to increase the age of aircraft.

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.

Separated flows and subsequent formation of shear layers are important fluid processes which play a dominant role in numerous engineering applications. Prediction of this fluid process is an important element in the design and analysis of highspeed vehicles and, ultimately, in the performance and trajectory analysis. The prediction methodology used in the current study includes the numerical solution of the Navier-Stokes equations on a multi-block grid system. Several turbulence models are used to investigate their performance for compressible, turbulent, separated and shear layer flows. The computational results are compared to available experimental data.

The architecture and training of artificial neural networks are briefly described. Five applications of these networks to design and analysis problems are presented; three in aerodynamics and two in flight dynamics. The aerodynamics cases are those of a harmonically oscillating airfoil, a pitching delta wing, and airfoil design. The flight dynamic examples involve control of a super maneuver and a decoupled control case. It is demonstrated that highly nonlinear aerodynamic cases can be generalized with sufficient accuracy for design purposes. It is shown that although neural networks generalize well on the aerodynamic problems, they appear lacking comparable robustness in modeling dynamic systems. It is also shown that generalization appears to become weak outside of the training domain.

The aerodynamic properties of a forward-swept wing were tested at low Reynolds numbers. The investigation was performed in a low-speed wind tunnel using a reflection plane model. Tunnel balance, model pressure taps, and flow visualization results were utilized to characterize the wing behavior over a range of Reynolds numbers from 0.25 × 106 - 0.75 × 106. In addition, the experimental data is compared to results obtained using a recently developed computer program known as WING3D. This modified Non-Planar Vortex Lattice Method program can calculate total wing lift and surface pressure distributions. The forward-swept wing has good aerodynamic qualities; in addition, the flow, on the outboard sections of the wing, remains attached beyond stall. The comparison of WING3D and experimental surface pressure distributions is good.