For the purpose of investigating briefly the effects of certain shape variables other than those discussed in the main body of the report, 10 supplementary airfoils were tested. The airfoil sections were as follows: 6 symmetrical sections with modified nose shapes, 2 sections with reflexed mean lines, and 2 sections simulating those of a wing having a flexible trailing edge.
Airfoils with modified nose shapes.—The airfoils of the first supplementary group investigated were developed from three of the symmetrical N.A.C.A. family airfoils: The N.A.C.A. 0006, the N.A.C.A. 0012, and the N.A.C.A. 0018. For each of these basic (or normal) sections one thinner-nosed section, denoted by the suffix T, and one blunter-nosed section, denoted by the suffix B, were developed and tested. The derivation of each modified section was similar to that of the normal section and was accomplished by a systematic change in the equation that defines the normal section. This change is principally a change in the nose radius, but it also results in modifications to the profile throughout its length, except at the maximum ordinate and at the trailing edge The nose radii of the sections in percent of the chord are as follows:
The aerodynamic characteristics of the modified sections are given in figures 72 to 77. These may be compared with the characteristics of the normal sections given in figures 4, 6, and 8. The maximum lift coefficients of the modified and the normal sections are plotted against the leading-edge radii in figure 97. It is interesting to note that the leading-edge radius is very critical in its effect on the maximum lift when the radius is small. This critical effect is also indicated by the rapid increase in the maximum lift with increasing thickness for the thin sections as shown in figure 85.
Airfoils with reflexed mean lines.—Previous investigations have shown that the pitching moment of cambered airfoils can be reduced by altering the form of the mean line toward the trailing edge, with a consequent loss of maximum lift but only a small reduction in drag. In order to compare the characteristics of sections of this type with those of the related sections of normal form, two airfoils were developed with the basic thickness distribution of the N.A.C.A. 0012 disposed about certain mean lines of the form given in reference 15
The values of h in this equation were chosen to give a camber of 0.02 and the values of (lambda) were chosen to give the airfoil designated the N.A.C.A. 2R212 a small negative moment and the airfoil designated the N.A.C.A. 2R212 a small positive moment. Characteristic curves for the two airfoils are given in figures 77 and 78. The principal characteristics of the sections may be conveniently compared with those of the related symmetrical section, the N.A.C.A. 0012, and a related normal section having a camber of 2 percent of the chord, the N.A.C.A. 2412, by means of the following table arranged in the order of increasing pitching-moment coefficients.
These results indicate that airfoils having reflexed mean lines may be of questionable value because of the adverse effect of this mean-line shape on the maximum lift coefficient.
Thickness and camber modifications near the trailing edge.—Two airfoils were developed to simulate an airfoil having a flexible trailing edge in a straight and in a given deflected position. The thickness distribution is composed of three parts: the forward portion (0 to 0.3c) having the same distribution as the N.A.C.A. 0012, the rear portion (from 0.7c to the trailing edge) having a thin, uniform value, and the central portion joining these two with fair curves. As shown in figure 80, the two airfoils differ only in the rear portion, the section designated N.A.C.A. 0012Fo simulating that of a wing having the trailing edge deformed for the high-speed condition, and the section designated N.A.C.A. 0012F1 simulating that of the same wing with the trailing edge bent down in a circular arc. Curves of the aerodynamic characteristics for both conditions are compared in figure 80. Considering the results given by both airfoils as two conditions for one airfoil, a very high maximum lift with a reasonably low minimum drag is obtained. On this assumption the ratio
is 197, slightly higher than the value of this ratio given by the N.A.C.A. 2412.
In order to study the effects of an extreme change in the thickness distribution, the principal characteristics of the two sections may be compared with those of the related normal sections, the N.A.C.A. 0012 and the N.A.C.A. 6712. The maximum lift coefficient is little affected by the change in the thickness distribution, but it is of interest to note (table I) that the slope of the lift curve of the N.A.C.A. 0012Fo is slightly greater than 2(pi) per radian, as compared with an appreciably lower slope for the N.A.C.A. 0012. The profile drag is also affected by the change in the thickness distribution. Of the two symmetrical sections, the profile drag of the N.A.C.A. 0012Fo is much higher than that of the N.A.C.A. 0012 over the entire lift range. This is not true, however, for the two cambered sections. Comparing the characteristics of the N.A.C.A. 0012F1 with those of the N.A.C.A. 6712, we find that at low values of the lift the profile drag of the former is much higher, but as the lift increases this difference becomes less, and in the high-lift range the profile drag of the N.A.C.A. 0012F, is considerably less than that of the N.A.C.A. 6712.
Table of Contents | Summary | Introduction | Description of Airfoils | Apparatus and Methods | Results | Discussion | Supplementary Airfoils | Conclusions | Appendix | References