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At least 1,189 records · Page 66
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

Determination of the Pressure Drag of Airfoils by Integration of Surface Pressures

A study was conducted of the causes of pressure drag of subsonic airfoils. In a previous paper by the author, the pressure drag is obtained by calculating the total drag from the momentum defect in the boundary layer at the trailing edge and subtracting the friction drag obtained from integration of surface friction along the chord. Herein, the pressure drag is obtained by integrating the streamwise components of surface pressure around the airfoil. Studies were made to verify the accuracy of the integration procedure. The values of pressure drag were much smaller than those obtained by the previous method. This lack of agreement is attributed to the difficulty of calculating boundary layer conditions in the vicinity of the trailing edge and to the extreme sensitivity of the circulation and lift to the trailing edge conditions. The results of these studies are compared with those of previous investigations.

Phillips, William H.↗
NASA NTRS1990

A numerical study of the effects of wind tunnel wall proximity on an airfoil model

A procedure was developed for modeling wind tunnel flows using computational fluid dynamics. Using this method, a numerical study was undertaken to explore the effects of solid wind tunnel wall proximity and Reynolds number on a two-dimensional airfoil model at low speed. Wind tunnel walls are located at varying wind tunnel height to airfoil chord ratios and the results are compared with freestream flow in the absence of wind tunnel walls. Discrepancies between the constrained and unconstrained flows can be attributed to the presence of the walls. Results are for a Mach Number of 0.25 at angles of attack through stall. A typical wind tunnel Reynolds number of 1,200,000 and full-scale flight Reynolds number of 6,000,000 were investigated. At this low Mach number, wind tunnel wall corrections to Mach number and angle of attack are supported. Reynolds number effects are seen to be a consideration in wind tunnel testing and wall interference correction methods. An unstructured grid Navier-Stokes code is used with a Baldwin-Lomax turbulence model. The numerical method is described since unstructured flow solvers present several difficulties and fundamental differences from structured grid codes, especially in the area of turbulence modeling and grid generation.

Potsdam, Mark↗
NASA NTRS1991

Prediction of ice shapes and their effect on airfoil performance

Calculations of ice shapes and the resulting drag increases are presented for experimental data on a NACA 0012 airfoil. They were made with a combination of LEWICE and interactive boundary-layer codes for a wide range of conditions which include air speed and temperature, the droplet size and liquid water content of the cloud, and the angle of attack of the airfoil. In all cases, the calculated results account for the drag increase due to ice accretion and, in general, show good agreement.

Shin, Jaiwon↗
NASA NTRS1991

Transonic shock-induced dynamics of a flexible wing with a thick circular-arc airfoil

Transonic shock boundary layer oscillations occur on rigid models over a small range of Mach numbers on thick circular-arc airfoils. Extensive tests and analyses of this phenomena have been made in the past but essentially all of them were for rigid models. A simple flexible wing model with an 18 pct. circular arc airfoil was constructed and tested in the Langley Transonic Dynamics Tunnel to study the dynamic characteristics that a wing might have under these circumstances. In the region of shock boundary layer oscillations, buffeting of the first bending mode was obtained. This mode was well separated in frequency from the shock boundary layer oscillations. A limit cycle oscillation was also measured in a third bending like mode, involving wind vertical bending and splitter plate motion, which was in the frequency range of the shock boundary layer oscillations. Several model configurations were tested, and a few potential fixes were investigated.

Bennett, Robert M.↗
NASA NTRS1991

A study of residual interference effects in adaptive wall testing of a 7-inch chord CAST-10-2/DOA 2 airfoil model

The residual interferences present in the testing of a CAST10-2/DOA2 airfoil model in the Langley 0.3 m Transonic Cryogenic Tunnel (0.3 m TCT) is examined. The 7.1 inch chord airfoil model was tested in the 0.3 m TCT adaptive wall test section with a nominal 13 inch square cross section. The test data obtained on the same model during different tunnel entries showed different levels of residual interference. The present study shows that for valid comparison of the test data from adaptive wall tunnels, it is necessary to account for residual interferences, in particular the blockage correction to the Mach number.

Murthy, A. V.↗
NASA NTRS1990

Acoustic radiation from a thin airfoil in nonuniform subsonic flows

Noise resulting from the interaction of a three dimensional gust is modeled directly from unsteady aerodynamic codes. This paper provides bench mark results for the case of a thin airfoil for comparison with the more complex case of a loaded airfoil. It is shown that the acoustic pressure pattern strongly depends on the value of a certain reduced frequency. The effects of an oblique gust are examined and shown to significantly reduce the acoustic power for certain gust parameters range. The Acoustic power versus the gust frequency exhibits a maximum. High frequency and non-compact source effects are also investigated.

Atassi, H. M.↗
NASA NTRS1991

Mach number validation of a new zonal CFD method (ZAP2D) for airfoil simulations

A closed-loop overlapped velocity coupling procedure has been utilized to combine a two-dimensional potential-flow panel code and a Navier-Stokes code. The fully coupled two-zone code (ZAP2D) has been used to compute the flow past a NACA 0012 airfoil at Mach numbers ranging from 0.3 to 0.84 near the two-dimensional airfoil C(lmax) point for a Reynolds number of 3 million. For these cases, the grid domain size can be reduced to 3 chord lengths with less than 3-percent loss in accuracy for freestream Mach numbers through 0.8. Earlier validation work with ZAP2D has demonstrated a reduction in the required Navier-Stokes computation time by a factor of 4 for subsonic Mach numbers. For this more challenging condition of high lift and Mach number, the saving in CPU time is reduced to a factor of 2.

Strash, Daniel J.↗
NASA NTRS1989

An airfoil pitch apparatus-modeling and control design

The study of dynamic stall of rapidly pitching airfoils is being conducted at NASA Ames Research Center. Understanding this physical phenomenon will aid in improving the maneuverability of fighter aircraft as well as civilian aircraft. A wind tunnel device which can linearly pitch and control an airfoil with rapid dynamic reponse is needed for such tests. To develop a mechanism capable of high accelerations, an accurate model and control system is created. The model contains mathematical representations of the mechanical system, including mass, spring, and damping characteristics for each structural element, as well as coulomb friction and servovalve saturation. Electrical components, both digital and analog, linear and nonlinear, are simulated. The implementation of such a high-performance system requires detailed control design as well as state-of-the-art components. This paper describes the system model, states the system requirements, and presents results of its theoretical performance which maximizes the structural and hydraulic aspects of this system.

Andrews, Daniel R.↗
NASA NTRS1991

Prediction of ice shapes and their effect on airfoil performance

Calculations of ice shapes and the resulting drag increases are presented for experimental data on an NACA 0012 airfoil. They were made with a combination of LEWICE and interactive boundary-layer codes for a wide range of conditions which include airspeed and temperature, the droplet size and liquid water content of the cloud, and the angle of attack of the airfoil. In all cases the calculated results account for the drag increase due to ice accretion and, in general, show good agreement with data.

Shin, Jaiwon↗
NASA NTRS1991

Airfoil transition and separation studies using an infrared imaging system

An infrared imaging system was used to detect the thermal signature of boundary-layer flow regimes on a NACA 0012 airfoil from zero angle of attack up to separation. The boundary-layer transition from laminar to turbulent flow and the onset of separation could be seen on the airfoil thermograms. The findings were compared against the behavior of aluminum foil tufts observable both visually and with the infrared imaging system. This arrangement offers the option of using the infrared imaging system both for flow regime detection through surface thermography and flow visualization by the aluminum foil tufts. Ultimately the surface temperature changes due to variation in the angle of attack of a lifting surface provide a means for interpretation of the boundary-layer flow regimes.

Gartenberg, Ehud↗
NASA NTRS1991

A quantitative study of unsteady compressible flow on an oscillating airfoil

Detailed interferometric measurements of the flow near the leading edge of an oscillating airfoil offer the first detailed experimental quantification of the locally compressible flow field that surrounds an oscillating airfoil at moderate subsonic Mach numbers. Interferograms obtained by a specially adapted real-time point-diffraction interferometry technique have revealed significant characteristics of this complex, and very rapidly varying, locally supersonic flow. Instantaneous pressure distributions determined from these interferograms document the effect of unsteadiness on the leading-edge flow environment.

Carr, L. W.↗
NASA NTRS1991

Aerodynamics of airfoils with vortex trapped by two spanwise fences

An airfoil configuration is presented whose lift is enhanced by a trapped-vortex flow field. Based on the research to date, it is recommended that two spanwise fences be used to enclose the trapped vortex and that the fence heights be adjusted so that the equilibrium condition can be achieved with little or no mass removal from the core region of the vortex. It is also shown that the vortex bubble can be located fore and aft on the airfoil to control aerodynamic parameters, such as the pitching moment. Applications of the high-lift concept presented here are briefly discussed.

Rossow, Vernon J.↗
NASA NTRS1991

Evaluation of tunnel sidewall boundary-layer-control systems for high-lift airfoil testing

An experimental study was conducted in the NASA Langley Low-Turbulence Pressure Tunnel to evaluate a suction sidewall boundary-layer-control (BLC) technique used in testing 2D high-lift airfoils. Sidewall BLC is required to maintain spanwise two-dimensionality of the flow over the airfoil at large angles of attack. A supercritical-type high-lift air-foil, equipped with a double-slotted flap and a leading-edge slat, was used for the study which was conducted at a Mach number of 0.2 and Reynolds numbers based on chord of 9 and 16 million. The sidewall BLC technique, which features distributed suction through porous endplates connected to a venting system, was able to control sidewall boundary-layer separation and maintain two-dimensional flow over the high-lift configuration for both Reynolds numbers tested. Discussions on porous endplate optimization and effects of suction on section lift are presented. Results obtained with the suction system were also compared with previous data obtained with a tangential blowing BLC system for the same high-lift configuration.

Paschal, K.↗
NASA NTRS1991

A unified viscous theory of lift and drag of 2-D thin airfoils and 3-D thin wings

A unified viscous theory of 2-D thin airfoils and 3-D thin wings is developed with numerical examples. The viscous theory of the load distribution is unique and tends to the classical inviscid result with Kutta condition in the high Reynolds number limit. A new theory of 2-D section induced drag is introduced with specific applications to three cases of interest: (1) constant angle of attack; (2) parabolic camber; and (3) a flapped airfoil. The first case is also extended to a profiled leading edge foil. The well-known drag due to absence of leading edge suction is derived from the viscous theory. It is independent of Reynolds number for zero thickness and varies inversely with the square root of the Reynolds number based on the leading edge radius for profiled sections. The role of turbulence in the section induced drag problem is discussed. A theory of minimum section induced drag is derived and applied. For low Reynolds number the minimum drag load tends to the constant angle of attack solution and for high Reynolds number to an approximation of the parabolic camber solution. The parabolic camber section induced drag is about 4 percent greater than the ideal minimum at high Reynolds number. Two new concepts, the viscous induced drag angle and the viscous induced separation potential are introduced. The separation potential is calculated for three 2-D cases and for a 3-D rectangular wing. The potential is calculated with input from a standard doublet lattice wing code without recourse to any boundary layer calculations. Separation is indicated in regions where it is observed experimentally. The classical induced drag is recovered in the 3-D high Reynolds number limit with an additional contribution that is Reynold number dependent. The 3-D viscous theory of minimum induced drag yields an equation for the optimal spanwise and chordwise load distribution. The design of optimal wing tip planforms and camber distributions is possible with the viscous 3-D wing theory.

Yates, John E.↗
NASA NTRS1992

Prediction of ice accretion on a swept NACA 0012 airfoil and comparisons to flight test results

In the winter of 1989-90, an icing research flight project was conducted to obtain swept wing ice accretion data. Utilizing the NASA Lewis Research Center's DHC-6 DeHavilland Twin Otter aircraft, research flights were made into known icing conditions in Northeastern Ohio. The icing cloud environment and aircraft flight data were measured and recorded by an onboard data acquisition system. Upon entry into the icing environment, a 24 inch span, 15 inch chord NACA 0012 airfoil was extended from the aircraft and set to the desired sweep angle. After the growth of a well defined ice shape, the airfoil was retracted into the aircraft cabin for ice shape documentation. The ice accretions were recorded by ice tracings and photographs. Ice accretions were mostly of the glaze type and exhibited scalloping. The ice was accreted at sweep angles of 0, 30, and 45 degrees. A 3-D ice accretion prediction code was used to predict ice profiles for five selected flight test runs, which include sweep angle of zero, 30, and 45 degrees. The code's roughness input parameter was adjusted for best agreement. A simple procedure was added to the code to account for 3-D ice scalloping effects. The predicted ice profiles are compared to their respective flight test counterparts. This is the first attempt to predict ice profiles on swept wings with significant scalloped ice formations.

Reehorst, Andrew L.↗
NASA NTRS1991

Experimental ice shape and performance characteristics for a multi-element airfoil in the NASA Lewis Icing Research Tunnel

A study of the ice accretion patterns and performance of characteristics of a multi-element airfoil was undertaken at the NASA-Lewis Icing Research Tunnel. Several configurations were examined to determine the ice shape and performance characteristics. The testing included glaze, rime, and mixed icing regimes. Tunnel cloud conditions were set to correspond to those typical of the operating environment for commercial transport aircraft. Measurements acquired included ice profile tracings and aerodynamic forces both during the accretion process and in a post-accretion evaluation over a range of angle of attack. Substantial ice accretions developed on the main wing, flaps, and slat surfaces. Force measurements indicate severe performance degradation, especially near CL max, for both light and heavy ice accretion. Frost was seen on the lower surface of the airfoil which was found to contribute significantly to the force components.

Berkowitz, Brian M.↗
NASA NTRS1992

Semi-empirical model for prediction of unsteady forces on an airfoil with application to flutter

A semi-empirical model is described for predicting unsteady aerodynamic forces on arbitrary airfoils under mildly stalled and unstalled conditions. Aerodynamic forces are modeled using second order ordinary differential equations for lift and moment with airfoil motion as the input. This model is simultaneously integrated with structural dynamics equations to determine flutter characteristics for a two degrees-of-freedom system. Results for a number of cases are presented to demonstrate the suitability of this model to predict flutter. Comparison is made to the flutter characteristics determined by a Navier-Stokes solver and also the classical incompressible potential flow theory.

Mahajan, Aparajit J.↗
NASA NTRS1992

The NASA Langley laminar-flow-control experiment on a swept, supercritical airfoil: Evaluation of initial perforated configuration

The initial evaluation of a large-chord, swept, supercritical airfoil incorporating an active laminar-flow-control (LFC) suction system with a perforated upper surface is documented in a chronological manner, and the deficiencies in the suction capability of the perforated panels as designed are described. The experiment was conducted in the Langley 8-Foot Transonic Pressure Tunnel. Also included is an evaluation of the influence of the proximity of the tunnel liner to the upper surface of the airfoil pressure distribution.

Harris, Charles D.↗