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REVIEW OF BASIC AERODYNAMIC PRINCIPLES                 15
     Most of the NACA airfoils are classified among three types of airfoils: the four
digit the five digit, and the series 6 sections. The nomenclature and meaning of
various digits are explained with the help of following examples:
1)  Four digit series, example NACA 2412
      2: The maximum camber of the meanline is 0.02c
      4: The position of the maximum camber is at 0.4c.
     12: The maximum thickness is 0.12c.
2)  Five digit series, example NACA 23012
      2: The maximum camber of the meanline is approximately equal to 0.02c.
       The design lift coefficient is 0.15 times the first digit of the series, which in
     this case is 2.
        30: The position of maximum camber is 0.30l2 = 0.15c.
      12: The maximum thickness is 0.12c.
3)  Series 6 sections, example NACA 653-418
    6: Series designation.
      5: The minimum pressure is at 0.5c.
       3: The drag coefficientis near mirtimum value over a range oflift coefficients
      of 0,3 above and below the design lift coefficient.
      4: The design lift coefficient is 0.4.
      18: The maximum thickness is 0.18c.
   The concept of design lift coefficient is illustrated in Fig. 1.16. It is the mean
value of the range of lift coefficients for which the sectional drag coefficient is
minimum.
For additional information on airfoil data, refer to Refs. 2 and 3.
                  Cl
Fig*1.16    Concept of desigrt lift coeffiaenL
16               PERFORMANCE, STABfLfTY, DYNAMICS, AND CONTROL
I-4.2  Wing Planform Parameters
      General planform parameters that are usefulin estimating aerodynamic data are
the planform area S, aspect ratio A, mean aerodynamic chord c, and the spanwise
location of the mean aerodynamic chord Ymac  These parameters are given by the
following expressions:
                                                      S = 2 [,' c(y) dy                                        (1.19)
                                   A = bS                            (1.20)
                                   - = ~[,'C2(y)dy  .                        (1.21)
                             2
                                              ymac = i= [' -(Y)Y dy                                 (1.22)
where c(y) is thelocal chord, y is the spanwise coordinate, and b is the wing span.
    For a corrventional straight wing (Fig. 1.17), Eqs. (1.19-1.22) reduce to
                                                             S = ~i;Cr (1 +A)                                             (1.23)
　
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