text: Figure 3.13 shows the flow patterns for 
  two basic supersonic airfoil sections and pro- 
  vides the approximate equations for lift,drag, 
  and lift curve slope. Since the wave drag is 
  the only factor of difference between -the two 
  airfoil sections, notice the configuration fac- 
  tors which affect the wave drag. For the 
  same thickness ratio, 
  the circular arc airfoil 
  would have a larger wedge angle formed 
  between the upper and lower surfaces at the 
  leading edge. At the same flight Mach num- 
  ber the larger angle at the leading edge would 
  form the stronger shock wave at the nose and 
  cause a greater pressure change on the circular 
  arc airfoil. 
  This same principle applies when 
  investigating 
  thickness. 
  the effect of airfoil 
  the wave drag coefficients for 
  that 
  Notice 
  both airfoils vary as the SQUARE of the 
  thickness ratio, e.g., if 
  the thickness ratio 
  were doubled, the wave drag coefhcient would 
  he four times as great. 
  If the thickness were 
  increased, the airflow at the leading edge will 
  experience a greater change in direction and 
  a stronger shock wave will be formed. This 
  powerful variation of wave drag with 
  thick- 
  ness ratio necessitates the use of very thin air- 
  foils with sharp leading edges for supersonic 
  flight. An additional consideration is that 
  thin airfoil sections favor the use of low aspect 
  ratios and high taper to obtain lightweight 
  structures and preserve stiffness and rigidity. 
  the 
  denominator of each of the equations for the 
  aerodynamic coefficients and indicates a de- 
  crease in each of these coefficients with an 
  increase in Mach number. Essentially, this 
  means that any aerodynamic surface becomes 
  less sensitive to changes in angle of attack at 
  higher Mach numbers. The decrease in lift 
  curve slope with Mach number has tremendous 
  the stability and control of 
  implications 
  high speed aircraft. The vertical tail becomes 
  less sensitive to angles of sideslip and the 
  directional stability of the aircraft will deteri- 
  orate with Mach number. The horizontal 
  tail of 
  
  appears in 