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NACA0006 (JF, WT, XF, NF)
NACA0008 (JF, XF, NF)
NACA0009 (JF, WT, XF, NF)
NACA0010 (JF, XF, NF)
NACA0012 (JF, WT, CFD, XF, NF)
NACA0013 (JF, XF, NF)
NACA0015 (JF, XF, NF)
NACA0016 (JF, XF, NF)
NACA0017 (JF, XF, NF)
NACA0018 (JF, XF, NF)
NACA0021 (JF, XF, NF)
NACA0024 (JF, XF, NF)
NACA0025 (JF, XF, NF)
NACA0030 (JF, XF, NF)
NACA1408 (JF, WT, XF, NF)
NACA1409 (JF, XF, NF)
NACA1410 (JF, WT, XF, NF)
NACA1412 (JF, WT, XF, NF)
NACA1415 (JF, XF, NF)
NACA1418 (JF, XF, NF)
NACA1420 (JF, XF, NF)
NACA1421 (JF, XF, NF)
NACA1424 (JF, XF, NF)
NACA2.4511 (JF, WT, XF, NF)
NACA2206 (JF, XF, NF)
NACA2209 (JF, XF, NF)
NACA2210 (JF, XF, NF)
NACA2212 (JF, XF, NF)
NACA2213 (JF, XF, NF)
NACA2215 (JF, XF, NF)
NACA2216 (JF, XF, NF)
NACA2218 (JF, XF, NF)
NACA2219 (JF, XF, NF)
NACA2221 (JF, XF, NF)
NACA2224 (JF, XF, NF)
NACA2309 (JF, XF, NF)
NACA2310 (JF, XF, NF)
NACA2312 (JF, XF, NF)
NACA2315 (JF, XF, NF)
NACA2317 (JF, XF, NF)
NACA2318 (JF, XF, NF)
NACA2321 (JF, XF, NF)
NACA2324 (JF, XF, NF)
NACA2408 (JF, XF, NF)
NACA2409 (JF, XF, NF)
NACA2410 (JF, XF, NF)
NACA2411 (JF, XF, NF)
NACA2412 (JF, WT, XF, NF)
NACA2414 (JF, WT, XF, NF)
NACA2415 (JF, WT, XF, NF)
NACA2416 (JF, XF, NF)
NACA2417 (JF, XF, NF)
NACA2418 (JF, WT, XF, NF)
NACA2421 (JF, WT, XF, NF)
NACA2424 (JF, WT, XF, NF)
NACA3409 (JF, XF, NF)
NACA3410 (JF, XF, NF)
NACA3412 (JF, XF, NF)
NACA3413 (JF, XF, NF)
NACA3414 (JF, XF, NF)
NACA3415 (JF, XF, NF)
NACA3418 (JF, XF, NF)
NACA3421 (JF, XF, NF)
NACA4312 (JF, XF, NF)
NACA4315 (JF, XF, NF)
NACA4318 (JF, XF, NF)
NACA4321 (JF, XF, NF)
NACA4324 (JF, XF, NF)
NACA4409 (JF, XF, NF)
NACA4410 (JF, XF, NF)
NACA4412 (JF, WT, XF, NF)
NACA4413 (JF, XF, NF)
NACA4415 (JF, WT, XF, NF)
NACA4416 (JF, XF, NF)
NACA4418 (JF, WT, XF, NF)
NACA4421 (JF, WT, XF, NF)
NACA4424 (JF, WT, XF, NF)
NACA5509 (JF, XF, NF)
NACA5510 (JF, XF, NF)
NACA5512 (JF, XF, NF)
NACA5514 (JF, XF, NF)
NACA5515 (JF, XF, NF)
NACA5518 (JF, XF, NF)
NACA5521 (JF, XF, NF)
NACA5524 (JF, XF, NF)
NACA6209 (JF, XF, NF)
NACA6210 (JF, XF, NF)
NACA6212 (JF, XF, NF)
NACA6215 (JF, XF, NF)
NACA6218 (JF, XF, NF)
NACA6221 (JF, XF, NF)
NACA6224 (JF, XF, NF)
NACA6409 (JF, WT, XF, NF)
NACA6410 (JF, XF, NF)
NACA6412 (JF, XF, NF)
NACA6415 (JF, XF, NF)
NACA6418 (JF, XF, NF)
NACA6421 (JF, XF, NF)
NACA6424 (JF, XF, NF)




NACA67-716 a=0.0


NACA67716a=0.0
New! Bulk download of polars in Excel format: NACA67-716%20a%3D0.0.xlsx


Info | DAT and DXF files | Shape | Venkataraman Fit | Applications (0) | Similarly shaped airfoils
Plain Flapped Shapes: 25% flap chord | 30% flap chord | 35% flap chord | 40% flap chord
Basic Data: cl vs α | cd vs α | cm vs α | cl/cd vs α | cd vs cl | cl/cd vs cl | cl vs cd | cm vs cl | Curve fits | Lift curve slopes
Data source comparison (Codes, CFD, Wind tunnel)
Compressibility Effects: cl vs α | cd vs α | cm vs α | cl/cd vs α | cd vs cl | cl/cd vs cl | cl vs cd | cm vs cl
Roughness Effects: cl vs α | cd vs α | cm vs α | cl/cd vs α | cd vs cl | cl/cd vs cl | cl vs cd | cm vs cl
Plain Flap Effects: cl,max vs flap deflection: all flap chords | 25% flap chord | 30% flap chord | 35% flap chord | 40% flap chord
Plain Flap Effects: cl vs α for various flap deflections: 25% flap chord | 30% flap chord | 35% flap chord | 40% flap chord
Plain Flap Effects: cm vs α for various flap deflections: 25% flap chord | 30% flap chord | 35% flap chord | 40% flap chord
Plain Flap Effects: cd vs α for various flap deflections: 25% flap chord | 30% flap chord | 35% flap chord | 40% flap chord
Disclaimer: All information presented herein is delivered without guarantee or warranty of any kind. The user assumes the entire risk of use of this information.

Curve fits of basic Javafoil data for NACA67-716 a=0.0, smooth airfoils, Mach 0:


Re: 500k | 1M | 3M | 6M

Re = 500k | 1M | 3M | 6M


Lift: Cl = f(α) = -2.304e-04*α3 + 7.962e-04*α2 + 1.234e-01*α + 4.211e-01
Drag: Cd = f(α) = -4.706e-06*α3 + 3.580e-04*α2 + -1.828e-03*α + 9.966e-03
Moment: Cm = f(α) = 5.355e-06*α3 + 4.441e-05*α2 + -3.880e-03*α + -5.527e-02
−20−15−10−505101520−0.500.511.5
Re 500k, JavaFoil DataRe 500k, JavaFoil Curve FitNACA67-716 a=0.0Cubic curve fit of JavaFoil data for smooth airfoil at Mach 0Section lift coefficient vs angle of attackCl = f(α) = -2.304e-04*α​3​ + 7.962e-04*α​2​ + 1.234e-01*α + 4.211e-01Angle of Attack (deg)Section Lift Coefficient c​l​

−20−15−10−50510152000.050.10.15
Re 500k, JavaFoil DataRe 500k, JavaFoil Curve FitNACA67-716 a=0.0Cubic curve fit of JavaFoil data for smooth airfoil at Mach 0Section drag coefficient vs angle of attackCd = f(α) = -4.706e-06*α​3​ + 3.580e-04*α​2​ + -1.828e-03*α + 9.966e-03Angle of Attack (deg)Section Drag Coefficient c​d​

−20−15−10−505101520−0.08−0.06−0.04−0.020
Re 500k, JavaFoil DataRe 500k, JavaFoil Curve FitNACA67-716 a=0.0Cubic curve fit of JavaFoil data for smooth airfoil at Mach 0Section moment coefficient vs angle of attackCm = f(α) = 5.355e-06*α​3​ + 4.441e-05*α​2​ + -3.880e-03*α + -5.527e-02Angle of Attack (deg)Section Moment Coefficient c​m​

Re = 1M | 500k | 3M | 6M


Lift: Cl = f(α) = -2.367e-04*α3 + 8.147e-04*α2 + 1.245e-01*α + 4.197e-01
Drag: Cd = f(α) = -9.177e-06*α3 + 3.551e-04*α2 + -1.160e-03*α + 6.095e-03
Moment: Cm = f(α) = 1.853e-06*α3 + 6.009e-05*α2 + -3.318e-03*α + -5.837e-02
−20−15−10−505101520−0.500.511.5
Re 1M, JavaFoil DataRe 1M, JavaFoil Curve FitNACA67-716 a=0.0Cubic curve fit of JavaFoil data for smooth airfoil at Mach 0Section lift coefficient vs angle of attackCl = f(α) = -2.367e-04*α​3​ + 8.147e-04*α​2​ + 1.245e-01*α + 4.197e-01Angle of Attack (deg)Section Lift Coefficient c​l​

−20−15−10−50510152000.050.10.150.2
Re 1M, JavaFoil DataRe 1M, JavaFoil Curve FitNACA67-716 a=0.0Cubic curve fit of JavaFoil data for smooth airfoil at Mach 0Section drag coefficient vs angle of attackCd = f(α) = -9.177e-06*α​3​ + 3.551e-04*α​2​ + -1.160e-03*α + 6.095e-03Angle of Attack (deg)Section Drag Coefficient c​d​

−20−15−10−505101520−0.1−0.08−0.06−0.04−0.020
Re 1M, JavaFoil DataRe 1M, JavaFoil Curve FitNACA67-716 a=0.0Cubic curve fit of JavaFoil data for smooth airfoil at Mach 0Section moment coefficient vs angle of attackCm = f(α) = 1.853e-06*α​3​ + 6.009e-05*α​2​ + -3.318e-03*α + -5.837e-02Angle of Attack (deg)Section Moment Coefficient c​m​

Re = 3M | 500k | 1M | 6M


Lift: Cl = f(α) = -2.464e-04*α3 + 8.528e-04*α2 + 1.263e-01*α + 4.159e-01
Drag: Cd = f(α) = -1.446e-05*α3 + 3.311e-04*α2 + -3.250e-04*α + 3.287e-03
Moment: Cm = f(α) = -1.739e-06*α3 + 6.401e-05*α2 + -2.788e-03*α + -6.026e-02
−20−15−10−505101520−0.500.511.5
Re 3M, JavaFoil DataRe 3M, JavaFoil Curve FitNACA67-716 a=0.0Cubic curve fit of JavaFoil data for smooth airfoil at Mach 0Section lift coefficient vs angle of attackCl = f(α) = -2.464e-04*α​3​ + 8.528e-04*α​2​ + 1.263e-01*α + 4.159e-01Angle of Attack (deg)Section Lift Coefficient c​l​

−20−15−10−50510152000.050.10.150.2
Re 3M, JavaFoil DataRe 3M, JavaFoil Curve FitNACA67-716 a=0.0Cubic curve fit of JavaFoil data for smooth airfoil at Mach 0Section drag coefficient vs angle of attackCd = f(α) = -1.446e-05*α​3​ + 3.311e-04*α​2​ + -3.250e-04*α + 3.287e-03Angle of Attack (deg)Section Drag Coefficient c​d​

−20−15−10−505101520−0.1−0.08−0.06−0.04−0.0200.02
Re 3M, JavaFoil DataRe 3M, JavaFoil Curve FitNACA67-716 a=0.0Cubic curve fit of JavaFoil data for smooth airfoil at Mach 0Section moment coefficient vs angle of attackCm = f(α) = -1.739e-06*α​3​ + 6.401e-05*α​2​ + -2.788e-03*α + -6.026e-02Angle of Attack (deg)Section Moment Coefficient c​m​

Re = 6M | 500k | 1M | 3M


Lift: Cl = f(α) = -2.503e-04*α3 + 8.643e-04*α2 + 1.271e-01*α + 4.146e-01
Drag: Cd = f(α) = -1.692e-05*α3 + 3.140e-04*α2 + 5.799e-05*α + 2.612e-03
Moment: Cm = f(α) = -2.530e-06*α3 + 6.238e-05*α2 + -2.635e-03*α + -6.078e-02
−20−15−10−505101520−0.500.511.5
Re 6M, JavaFoil DataRe 6M, JavaFoil Curve FitNACA67-716 a=0.0Cubic curve fit of JavaFoil data for smooth airfoil at Mach 0Section lift coefficient vs angle of attackCl = f(α) = -2.503e-04*α​3​ + 8.643e-04*α​2​ + 1.271e-01*α + 4.146e-01Angle of Attack (deg)Section Lift Coefficient c​l​

−20−15−10−50510152000.050.10.150.2
Re 6M, JavaFoil DataRe 6M, JavaFoil Curve FitNACA67-716 a=0.0Cubic curve fit of JavaFoil data for smooth airfoil at Mach 0Section drag coefficient vs angle of attackCd = f(α) = -1.692e-05*α​3​ + 3.140e-04*α​2​ + 5.799e-05*α + 2.612e-03Angle of Attack (deg)Section Drag Coefficient c​d​

−20−15−10−505101520−0.1−0.08−0.06−0.04−0.0200.02
Re 6M, JavaFoil DataRe 6M, JavaFoil Curve FitNACA67-716 a=0.0Cubic curve fit of JavaFoil data for smooth airfoil at Mach 0Section moment coefficient vs angle of attackCm = f(α) = -2.530e-06*α​3​ + 6.238e-05*α​2​ + -2.635e-03*α + -6.078e-02Angle of Attack (deg)Section Moment Coefficient c​m​