NAVWEPS OO-EOT-RO  BASIC AERODYNAMICS  40 percent chord.  When this section is com-  pared with the NACA 0006 section the effect  of camber can be appreciated. At low lift  coefficients the thtn, symmetrical section has  much lower drag.  However, at lift coeffi-  cients above 03 the thicker, cambered section  has the lower drag.  Thus, proper camber and  thickness can improve the lift-drag ratio of  the section.  weight if the airplane is flown at the angle  of attack for (L/D),.  Of course, the gross  weight would affect the glide airspeed neces-  sary for this particular angle of attack but the  glide ratio would be unaffected.  AIRFOIL  DRAG  CHARACTERISTICS.  The total drag of an airplane is composed of  the drags of the individual components and  the forces caused by interference  between these  components. The drag of an airplane con-  figuration must include the various drags due  to lift, form, friction, interference, leakage,  etc. To appreciate the factors which affect  the drag of an airplane configuration, it is  most logical to consider the factors which  affect the drag of airfoil sections.  In order to  allow an objective consideration of the effects  of thickness, camber, etc., the properties of  two-dimensional sections must be studied.  Airfoil section properties are derived from the  basic profile in two-dimensional. flow and are  provided the lower case shorthand notation  to distinguish them from wing or airplane  properties, e.g., wing or airplane drag coe5-  cient is C, while airfoil section drag coefficient  is c,.  The NACA 63,412 is a cambered 12 percent  thick airfoil of the ‘“laminar  flow”  type.  This airfoil is shaped to produce a design lift  coe5cient of 0.4. Notice that the drag curve  of this airfoil has distinct aberrations with  very low drag coefficients near the lift coeffi-  cient of 0.4.  This airfoil profile has its camber  and thickness distributed to produce very low  uniform velocity on the forward surface  (mini-  mum pressure  point well aft) at this lift coeffi-  cient.  The resulting pressure and velocity  distribution enhance extensive laminar flow  in the boundary layer and greatly reduce the  skin friction drag.  The benefit of the laminar  flow is appreciated by comparing the minimum  drag of this airfoil with an airfoil which has  one-half the maximum thickness-the NACA  ooo6.  The drag characteristics  of three illustrative  airfoil sections are shown in figure 1.14.  The  section drag coe&cient, c,, is plotted versus  the section lift coefficient, cr. The drag on  the airfoil section is composed of pressure  drag  and skin friction.  When the airfoil is at low  lift coe&cients, the drag due to skin friction  predominates. The drag curve for a conven-  tional airfoil tends to be quite shallow in this  region since there is very little variation of  skin friction with angle of attack.  When the  airfoil is at high lift coefficients, form or  pressure drag predominates and the drag co-  efficient varies rapidly with lift  coefficient.  The NACA 0006 is a thin symmetrical profile  which has a maximum thickness of 6 percent  located at 30 percent of the chord.  This  section shows a typical variation of cd and cr.  The choice of an airfoil section will depend  on the consideration oftmany different factors.  While the cI, of the section is an important  quality, a more appropriate factor for con-  sideration is the maximum lift coefficient of  the section when various high lift devices are  applied. Trailing edge flaps and leading edge  high lift devices are applied to increase the  cr,,  for low speed performance. Thus, an  appropriate factor for comparison is the ratio  of section drag coe5cient to section maximum  lift coefficient with flaps-cd/crm,. When this  quantity is corrected for compressibility, a  preliminary selection of an airfoil section is  possible. The airfoil having the lowest value  of c&~, at the design flight condition (en-  durance, range, high speed, etc.) will create  the least section drag for a given .design stall  speed.  The NACA 4412 section is a 12 percent thick  airfoil with 4 percent maximum camber at 