Compressibility effects on a hovering helicopter rotor having an NACA 0018 root airfoil tapering to an NACA 0012 tip airfoil
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Powell, Robert D , Jr↗
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CFD Study of NACA 0018 Airfoil with Flow Control
The abilities of two different Reynolds-Averaged Navier-Stokes codes to predict the effects of an active flow control device are evaluated. The flow control device consists of a blowing slot located on the upper surface of an NACA 0018 airfoil, near the leading edge. A second blowing slot present on the airfoil near mid-chord is not evaluated here. Experimental results from a wind tunnel test show that a slot blowing with high momentum coefficient will increase the lift of the airfoil (compared to no blowing) and delay flow separation. A slot with low momentum coefficient will decrease the lift and induce separation even at low angles of attack. Two codes, CFL3D and FUN3D, are used in two-dimensional computations along with several different turbulence models. Two of these produced reasonable results for this flow, when run fully turbulent. A more advanced transition model failed to predict reasonable results, but warrants further study using different inputs. Including inviscid upper and lower tunnel walls in the simulations was found to be important in obtaining pressure distributions and lift coefficients that best matched experimental data. A limited number of three-dimensional computations were also performed.
Eggert, Christopher A.↗
The Flow About a Section of a Finite-aspect-ratio NACA 0018 Airfoil on a Transonic Bump
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Mellenthin, Jack A↗
Effectiveness of Thermal-Pneumatic Airfoil-Ice-Protection System
Icing and drag investigations were conducted in the NACA Lewis icing research tunnel employing a combination thermal-pneumatic de-icer mounted on a 42-inch-chord NACA 0018 airfoil. The de-icer consisted of a 3-inch-wide electrically heated strip symmetrically located about the leading edge with inflatable tubes on the upper and lower airfoil surfaces aft of the heated area. The entire de-icer extended to approximately 25 percent of chord. A maximum power density of 9.25 watts per square inch was required for marginal ice protection on the airfoil leading edge at an air temperature of 00 F and an airspeed of 300 miles per hour. Drag measurements indicated, that without icing, the de-icer installation increased the section drag to approximately 140 percent of that of the bare airfoil; with the tubes inflated, this value increased to a maximum of approximately 620 percent. A 2-minute tube-inflation cycle prevented excessive ice formation on the inflatable area although small scattered residual Ice formations remained after inflation and were removed intermittently during later cycles. Effects of the time lag of heater temperatures after initial application of power and the insulating effect of ice formations on heater temperatures were also determined.
Gowan, William H., Jr.↗
A study of flow past an airfoil with a jet issuing from its lower surface
The aerodynamics of a NACA 0018 airfoil with a rectangular jet of finite aspect ratio exiting from its lower surface at 90 deg to the chord were investigated. The jet was located at 50% of the wing chord. Measurements include static pressures on the airfoil surface, total pressures in the near wake, and local velocity vectors in different planes of the wake. The effects of jet cross flow interaction on the aerodynamics of the airfoil are studied. It is indicated that at all values of momentum coefficients, the jet cross flow interaction produces a strong contra-rotating vortex structure in the near wake. The flow behind the jet forms a closed recirculation region which extends up to a chord length down stream of the trailing edge which results in the flow field to become highly three dimensional. The various aerodynamic force coefficients vary significantly along the span of the wing. The results are compared with a jet flap configuration.
Krothapalli, A.↗
Numerical simulation by TVD schemes of complex shock reflections from airfoils at high angle of attack
The shock-capturing capability of total variation diminishing (TVD) schemes is demonstrated for a more realistic complex shock-diffraction problem for which the experimental data are available. Second-order explicit upwind and symmetric TVD schemes are used to solve the time-dependent Euler equations of gas dynamics for the interaction of a blast wave with an airfoil at high angle-of-attack. The test cases considered are a time-dependent moving curved-shock wave and a contant moving planar-shock wave impinging at an angle-of-attack 30 deg on a NACA 0018 airfoil. Good agreement is obtained between isopycnic contours computed by the TVD schemes and those from experimental interferograms. No drastic difference in flow-field structure is found between the curved- and planar-shock wave cases, except for a difference in density level near the lower surface of the airfoil. Computation for cases with higher shock Mach numbers is also possible. Numerical experiments show that the symmetric TVD scheme is less sensitive to the boundary conditions treatment than the upwind scheme.
Moon, Young J.↗
Construction of explicit and implicit symmetric TVD schemes and their applications
A one-parameter family of explicit and implicit total variation diminishing (TVD) schemes is developed which permits incorporation of an expanded group of slope and flux limiters. The numerical technique is intended for use in calculations which include a time-differencing scheme and an optional Lax-Wendroff scheme. Methods of extending the TVD models to nonlinear scalar equations and systems of hyperbolic conservation equations are described. Sample results are presented from calculations of shocked flows around NACA 0012 and NACA 0018 airfoils.
Yee, H. C.↗
Use of Plasma Actuators as a Moving-Wake Generator
The work documented in this report tests the concept of using plasma actuators as a simple and easy way to generate a simulated moving-wake and the disturbances associated with it in turbines. This wake is caused by the blades of the upstream stages of the turbine. Two types of devices, one constructed of arrays of NACA 0018 airfoils, and the one constructed of flat plates were studied. The airfoils or plates were equipped with surface mounted dielectric barrier discharge (DBD) plasma actuators, which were used to generate flow disturbances resembling moving-wakes. CTA hot-wire anemometry and flow visualization using a smoke-wire were used to investigate the wake independence at various spacings and downstream locations. The flat plates were found to produce better results than the airfoils in creating large velocity fluctuations in the free-stream flow. Different dielectric materials, plasma actuator locations, leading edge contours, angles of attack and plate spacings were investigated, some with positive results. The magnitudes of the velocity fluctuations were found to be comparable to existing mechanical moving-wake generators, thus proving the feasibility of using plasma actuators as a moving-wake generator.
Corke, Thomas C.↗
Unsteady Thick Airfoil Aerodynamics: Experiments, Computation, and Theory
An experimental, computational and theoretical investigation was carried out to study the aerodynamic loads acting on a relatively thick NACA 0018 airfoil when subjected to pitching and surging, individually and synchronously. Both pre-stall and post-stall angles of attack were considered. Experiments were carried out in a dedicated unsteady wind tunnel, with large surge amplitudes, and airfoil loads were estimated by means of unsteady surface mounted pressure measurements. Theoretical predictions were based on Theodorsen's and Isaacs' results as well as on the relatively recent generalizations of van der Wall. Both two- and three-dimensional computations were performed on structured grids employing unsteady Reynolds-averaged Navier-Stokes (URANS). For pure surging at pre-stall angles of attack, the correspondence between experiments and theory was satisfactory; this served as a validation of Isaacs theory. Discrepancies were traced to dynamic trailing-edge separation, even at low angles of attack. Excellent correspondence was found between experiments and theory for airfoil pitching as well as combined pitching and surging; the latter appears to be the first clear validation of van der Wall's theoretical results. Although qualitatively similar to experiment at low angles of attack, two-dimensional URANS computations yielded notable errors in the unsteady load effects of pitching, surging and their synchronous combination. The main reason is believed to be that the URANS equations do not resolve wake vorticity (explicitly modeled in the theory) or the resulting rolled-up un- steady flow structures because high values of eddy viscosity tend to \smear" the wake. At post-stall angles, three-dimensional computations illustrated the importance of modeling the tunnel side walls.
Strangfeld, C.↗
Tests of NACA 0009, 0012, and 0018 Airfoils in the Full-Scale Tunnel
An investigation was conducted in the NACA full-scale wind tunnel to determine the aerodynamic characteristics of the NACA 0009, 0012, and 0018 airfoils, with the ultimate purpose of providing data to be used as a basis for comparison with other wind-tunnel data, mainly in the study of scale and turbulence effects. Three symmetrical 6 by 36-foot rectangular airfoils were used. The Reynolds number range for minimum drag was form 1,800,000 to 7,000,000 and for maximum lift, from 1,700,000 to 4,500,000. The effect of rounded tips was determined for each of the airfoils. Tests were also made of the NACA 0012 airfoil equipped with a 0.20c full-span split flap hinged at 0.80c. Tuft surveys were included to show the progressive breakdown of flow near maximum lift. Momentum surveys were made in conjunction with force measurements at zero lift as an aid in converting force-test data to section coefficients.
Goett, Harry J↗
A modification to thin-airfoil-section theory, applicable to arbitrary airfoil sections, to account for the effects of thickness on the lift distribution
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AIRFOILS, CAMBERED↗
Wind-Tunnel Tests of 10-foot-diameter Autogiro Rotors
Report presents the results of a series of 10-foot-diameter autogiro rotor models tested in the NACA 20-foot wind tunnel. Four of the models differed only in the airfoil sections of the blades, the sections used being the NACA 0012, 0018, 4412, and 4418. Three additional models employing the NACA 0012 section were tested, in which a varying portion of the blade near the hub was replaced by a streamline tube with a chord of about one-fourth the blade chord.
Wheatley, John B↗
Experimental Investigation of the Momentum Method for Determining Profile Drag
Report presents the results of an experimental investigation conducted in the full-scale tunnel to determine the accuracy of the Jones and the Betz equations for computing profile drag from total and static pressure surveys in the wake of wings. Surveys were made behind 6 by 8-foot airfoils of the NACA 0009, and 0018 sections at zero lift and behind the NACA 0012 at positive lifts. The surveys were made at various spanwise positions and at distances behind the airfoil ranging from 0.05c to 3.00c.
Goett, Harry J↗
Determination of Boundary-Layer Transition on Three Symmetrical Airfoils in the NACA Full-Scale Wind Tunnel
For the purpose of studying the transition from laminar to turbulent flow, boundary-layer measurements were made in the NACA full-scale wind tunnel on three symmetrical airfoils of NACA 0009, 0012, and 0018 sections. The effects of variations in lift coefficient, Reynolds number, and airfoil thickness on transition were investigated. Air speed in the boundary layer was measured by total-head tubes and by hot wires; a comparison of transition as indicated by the two techniques was obtained. The results indicate no unique value of Reynolds number for the transition, whether the Reynolds number is based upon the distance along the chord or upon the thickness of the boundary layer at the transition point. In general, the transition is not abrupt and occurs in a region that varies in length as a function of the test conditions.
Silverstein, Abe↗
Tests of the NACA 0025 and 0035 Airfoils in the Full-Scale Wind Tunnel
This report presents the results of an investigation conducted in the NACA full-scale wind tunnel to determine the aerodynamic characteristics of the 6 by 36-foot rectangular NACA 0025 and 0035 airfoils. The aerodynamic characteristics of the plain airfoils with rounded and square tips were determined by force tests through a complete angle-of-attack range, in addition, the profile drag was determined by the momentum method. The transition points on the airfoils were located by boundary-layer determinations with small total-head and static tubes. Each airfoil was also tested with a 0.20c full-span split flap. Tuft surveys were included to show the progressive breakdown of flow with increasing angles of attack. Previously published data from tests of the NACA 0009, 0012, and 0018 airfoils in the full-scale tunnel have been included in the summary curves.
Bullivant, W Kenneth↗