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NACA 5 digit airfoil generator (NACA24012 AIRFOIL)


Airfoil NACA 24012 Airfoil cl=0.30 T=12.0% P=20.0%

Dat file

NACA 24012 Airfoil cl=0.30 T=12.0% P=20.0%
  1.000034  0.001260
  0.998499  0.001517
  0.993901  0.002286
  0.986271  0.003554
  0.975654  0.005302
  0.962115  0.007503
  0.945737  0.010126
  0.926621  0.013135
  0.904884  0.016488
  0.880660  0.020143
  0.854096  0.024054
  0.825357  0.028176
  0.794619  0.032461
  0.762070  0.036859
  0.727912  0.041322
  0.692353  0.045797
  0.655613  0.050232
  0.617917  0.054570
  0.579496  0.058755
  0.540587  0.062726
  0.501429  0.066422
  0.462262  0.069781
  0.423325  0.072741
  0.384859  0.075241
  0.347100  0.077226
  0.310278  0.078646
  0.274563  0.079453
  0.239831  0.079489
  0.206317  0.078497
  0.174345  0.076299
  0.144248  0.072811
  0.116354  0.068044
  0.090970  0.062103
  0.068368  0.055171
  0.048771  0.047489
  0.032351  0.039329
  0.019219  0.030969
  0.009432  0.022669
  0.002997  0.014646
 -0.000123  0.007059
  0.000000  0.000000
  0.003205 -0.006286
  0.009315 -0.011612
  0.018198 -0.016059
  0.029725 -0.019740
  0.043770 -0.022790
  0.060222 -0.025349
  0.078992 -0.027560
  0.100013 -0.029550
  0.123240 -0.031426
  0.148645 -0.033265
  0.176207 -0.035103
  0.205898 -0.036930
  0.237671 -0.038675
  0.271447 -0.040202
  0.307039 -0.041310
  0.343883 -0.041880
  0.381695 -0.041935
  0.420240 -0.041514
  0.459279 -0.040660
  0.498571 -0.039420
  0.537872 -0.037842
  0.576938 -0.035976
  0.615529 -0.033871
  0.653404 -0.031573
  0.690330 -0.029128
  0.726079 -0.026578
  0.760428 -0.023965
  0.793166 -0.021330
  0.824091 -0.018710
  0.853011 -0.016145
  0.879746 -0.013673
  0.904133 -0.011331
  0.926019 -0.009155
  0.945270 -0.007183
  0.961765 -0.005448
  0.975403 -0.003980
  0.986099 -0.002808
  0.993787 -0.001953
  0.998419 -0.001434
  0.999966 -0.001260

NACA 5 digit airfoils in the database

NACA 22112 NACA 23012
NACA 23015 NACA 23018
NACA 23021 NACA 23024
NACA 23112 NACA 24112
NACA 25112
First digit = Cl * 20/3 (0.05 to 1)
Second & third digits
Fourth & fifth digits. (1 to 30%)
20 to 200
Cosine or linear spacing
Open or closed TE
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NACA 5 digit airfoil specification

The NACA 5 digit airfoils use the same thickness envelope as the 4 series but with a different camber line and numbering system.

NACA LPQXX
e.g.
NACA 23012

DigitsLetterExampleDescription
1L2This digit controls the camber. It indicates the designed coefficient of lift (Cl) multiplied by 3/20. In the examble L=2 so Cl=0.3
2P3The position of maximum camber divided by 20. In the examble P=3 so maximum camber is at 0.15 or 15% chord
3Q00 = normal camber line, 1 = reflex camber line
4 & 5XX12The maximum thickness as percentage.In the examble XX=12 so the maximum thickness is 0.12 or 12% chord.

NACA 5 digit airfoil calculation

The equation for the camber line is split into two sections like the 4 digit series but the division between the two sections is not at the point of maximum camber. There are also different equations for standard and reflex camber lines.

NACA 5 digit airfoil camber line calculation

The values for the constants r, k1 and k2/k1 are tabulated for various positions of the maximum camber at a coefficient of lift (Cl) value of 0.3. The camber and gradient can be scaled linearly to the required Cl value.

DescriptionDigitsCamber
position(%)
rk1k2/k1
5% standard1050.0580361.400
10% standard20100.126051.640
15% standard30150.202515.957
20% standard40200.29006.643
25% standard50250.39103.230
10% reflex21100.130051.9900.000764
15% reflex31150.217015.7930.00677
20% reflex41200.31806.5200.0303
25% reflex51250.44103.1910.1355

Having calculated the camber line, the thickness distribution, calculation of the airfoil envelope and plotting of coordinates is done in the same way as the naca 4 digit airfoils. Details can be found here

References