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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 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.

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.

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.

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.

The motivation for this project was to design, and develop an aircraft seat to meet the proposed FAA 32g vertical/longitudinal dynamic test requirements specified in NPRM 93-71. A major goal of the design was to develop a production-quality seat in terms of weight, comfort, appearance, simplicity, and manufacturability. The relevant injury criteria was to obtain an occupant lumbar (spinal) load below 6670 N (1500 1bf). The design incorporated energy absorbing devices in the cushion and chair legs. The seat developed was based on the Beech King Air design and incorporated a modified seat frame, seat back, and reclining mechanism. The seat cushions were provided by Oregon Aero, while the seat pan and seat legs were designed and manufactured at WSU.

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.

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.

Two new concept of flexible fixture subsystem (FFS) for aircraft wing spar assembly are introduced in this paper. The advantages and characteristics of FFS are discussed and compared with the current assembly method and fixtures. The objective of FFS is to replace the dedicated tooling and be able to quickly reconfigure itself for new types of spars. The fixture enables a family of spars to be mounted and assembled in the same tooling. Left- and right-hand side spars, varying lofts(spar cap angles), height, and depths are all accommodated on the same tool, within its envelop.

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 Federal Aviation Administration (FAA)'s analysis of commuter aircraft accidents and ongoing research has indicated that the crashworthiness capabilities of smaller aircraft may be questionable. The small size of these aircraft results in a stiff structure and consequently higher impact loads experienced by the occupants. In 1993, the FAA issued a Notice of Proposed Rule Making (NPRM) 93-71 to increase the deceleration pulse amplitude of the sled tests under the Test-1 conditions (60-degree test) to 32G for the commuter type aircraft. To meet this condition, the seat design must exploit the energy absorption potential for its structural components. Energy absorbing components may include the seat legs, seat pan, and seat cushion. The intent is to design the seat so that it strikes well beyond the elastic limit to absorb the energy of the impact. To date, no seat has yet been able to pass the proposed criteria with an acceptable limit on the lumbar load (1500 pounds).

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.