Problems associated with lateral-control devices, leading-edge air intakes, and interference are briefly discussed, together with aerodynamic problems of application. Available data on high-lift devices are presented. The report includes an analysis of the lift, drag, pitching-moment, and critical-speed characteristics of the airfoils, together with a discussion of the effects of surface conditions. Detail data necessary for the application of the airfoils to wing design are presented in supplementary figures placed at the end of the paper. The general methods used to derive the basic thickness forms for NACA 6 and 7-series airfoils together with their corresponding pressure distributions are presented. New data are presented that permit the rapid calculation of the approximate pressure distributions for the older NACA four-digit and five-digit airfoils by the same methods used for the NACA 6-series airfoils. Left click over the numeric value and a text box will open, type the desired value, close the window, and the value will be updated.Summary of Airfoil Data The historical development of NACA airfoils is briefly reviewed. Values used to define the thickness and camber line may also be entered manually. Camber line slider bars are only used for the four digit airfoils, five digit airfoil camber lines are the NACA default values. The slider bars are used to change the thickness and camber line parameters. 4 You can replace the equation above with any bounded function. Various modification of the mean line, leading edge radius of curvature, and trailing edge angle were made as a part of the NACA research however, these are not implemented in AeroFoil. Where x u and y u are the upper surface coordinates, x l and y l are the lower surface coordinates, and θ is the angle between the mean line and the airfoil chord. The mean line and thickness distribution are combined to give the airfoil shape by the following: Last two digits describing maximum thickness of the airfoil as percent of the chord. Second digit describing the distance of maximum camber from the airfoil leading edge in tenths of the chord. Values of k 1 were calculated to result in a design lift coefficient of 0.3. The NACA four-digit wing sections define the profile by: 2 First digit describing maximum camber as percentage of the chord. Values of m were chosen to give five positions of p, the maximum camber, of 0.05c, 0.10c, 0.15c, 0.20c, and 0.25c, where c in the airfoil chord. The historical development of NACA airfoils is briefly reviewed. The historical development of NACA airfoils is briefly reviewed. Y c = k 1 m 3 (1 - x) from x = m to x = 1 Airfoil database search (NACA 6 series) Search the 1638 airfoils available in the databases filtering by name, thickness and camber. The camber line for a 5-digit airfoil is given by: Where m is the maximum ordinate of the mean line expressed as a fraction of the chord and p is the chordwise location of the maximum ordinate. Y c = m / (1 - p) 2 aft of the maximum ordinate The early NACA airfoil series, the 4-digit, 5-digit, and modified 4-/5-digit, were generated using analytical equations that describe the camber (curvature) of the mean-line (geometric centerline) of the airfoil section as well as the section's thickness distribution along the length of the airfoil. Y c = m (2 p x - x 2) / p 2 forward of the maximum ordinate, and The camber line for a 4-digit airfoil is given by: Modified Naca 0012 airfoil for Sandia VAWT tests. Where t is the thickness expressed as a fraction of the chord. NACA 23112 5 digit reflex airfoil Max thickness 12 at 29.5 chord Max camber 1.2 at 14.7 chord Source Javafoil generated (naca24112-jf) NACA 24112: Airfoil details Send to airfoil plotter Add to comparison Lednicer format dat file Selig format dat file: NACA 24112 5 digit reflex airfoil Max thickness 12 at 29.4 chord Max camber 1.7 at 20. (naca0012h-sa) NACA0012H for VAWT from Sandia report SAND80-2114. The available data was used to create an equation for a thickness distribution that was similar in turn to those airfoils. In the early 30’s, NACA found that the successful airfoils of the time had very similar thickness distributions after their camber lines were straightened. NACA 4 and 5 digit wing sections combine a thickness distribution with a camber line. AeroFoil Help NACA 4 and 5 digit Airfoils
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