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NACA 5 digit airfoil generator (NACA15110 AIRFOIL)

Dat file

NACA 15110 Airfoil cl=0.20 T=10.0% P=25.0%
  0.999993  0.001050
  0.998451  0.001221
  0.993835  0.001732
  0.986174  0.002581
  0.975519  0.003764
  0.961938  0.005275
  0.945519  0.007106
  0.926365  0.009249
  0.904599  0.011692
  0.880358  0.014421
  0.853792  0.017418
  0.825068  0.020664
  0.794360  0.024133
  0.761858  0.027795
  0.727759  0.031616
  0.692271  0.035554
  0.655607  0.039561
  0.617988  0.043583
  0.579641  0.047558
  0.540797  0.051421
  0.501688  0.055103
  0.462553  0.058530
  0.423611  0.061629
  0.384998  0.064280
  0.346935  0.066319
  0.309667  0.067601
  0.273443  0.068007
  0.238519  0.067455
  0.205151  0.065902
  0.173592  0.063349
  0.144084  0.059839
  0.116857  0.055452
  0.092121  0.050306
  0.070060  0.044544
  0.050833  0.038326
  0.034568  0.031820
  0.021360  0.025189
  0.011276  0.018582
  0.004353  0.012124
  0.000599  0.005911
  0.000000  0.000000
  0.002483 -0.005395
  0.007959 -0.010091
  0.016354 -0.014116
  0.027584 -0.017515
  0.041553 -0.020347
  0.058160 -0.022680
  0.077299 -0.024587
  0.098862 -0.026146
  0.122737 -0.027430
  0.148809 -0.028507
  0.176960 -0.029438
  0.207064 -0.030269
  0.238983 -0.031031
  0.272567 -0.031736
  0.307650 -0.032379
  0.344048 -0.032930
  0.381556 -0.033342
  0.419955 -0.033547
  0.458988 -0.033468
  0.498312 -0.033066
  0.537662 -0.032358
  0.576793 -0.031362
  0.615457 -0.030102
  0.653410 -0.028599
  0.690412 -0.026878
  0.726231 -0.024967
  0.760641 -0.022896
  0.793425 -0.020700
  0.824381 -0.018417
  0.853314 -0.016091
  0.880048 -0.013768
  0.904418 -0.011498
  0.926275 -0.009332
  0.945488 -0.007323
  0.961941 -0.005521
  0.975537 -0.003974
  0.986196 -0.002723
  0.993853 -0.001802
  0.998466 -0.001239
  1.000007 -0.001050

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

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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

Digits	Letter	Example	Description
1	L	2	This 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
2	P	3	The position of maximum camber divided by 20. In the examble P=3 so maximum camber is at 0.15 or 15% chord
3	Q	0	0 = normal camber line, 1 = reflex camber line
4 & 5	XX	12	The 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.

Description	Digits	Camber position(%)	r	k₁	k₂/k₁
5% standard	10	5	0.0580	361.400
10% standard	20	10	0.1260	51.640
15% standard	30	15	0.2025	15.957
20% standard	40	20	0.2900	6.643
25% standard	50	25	0.3910	3.230
10% reflex	21	10	0.1300	51.990	0.000764
15% reflex	31	15	0.2170	15.793	0.00677
20% reflex	41	20	0.3180	6.520	0.0303
25% reflex	51	25	0.4410	3.191	0.1355

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

NACA 5 digit airfoil generator (NACA15110 AIRFOIL)

Dat file

NACA 5 digit airfoil specification

NACA 5 digit airfoil calculation

References