Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

AH 85-L-120/17 (ah85l120-il)

AH 85-L-120/17 - Althaus AH 85-L-120/17 symmetrical airfoil


Airfoil ah85l120-il
Details Dat file Parser  
(ah85l120-il) AH 85-L-120/17
Althaus AH 85-L-120/17 symmetrical airfoil
Max thickness 12.1% at 43.5% chord.
Max camber 0% at 0% chord
Source UIUC Airfoil Coordinates Database
Source dat file
The dat file is in Selig format
AH 85-L-120/17
1.00000     0.00000
0.99893     0.00080
0.99572     0.00112
0.99039     0.00167
0.98296     0.00229
0.97347     0.00314
0.96194     0.00422
0.94844     0.00564
0.93301     0.00746
0.91573     0.00976
0.89668     0.01260
0.87592     0.01602
0.85355     0.02003
0.82967     0.02455
0.80438     0.02941
0.77779     0.03434
0.75000     0.03903
0.72114     0.04330
0.69134     0.04707
0.66072     0.05032
0.62941     0.05308
0.59755     0.05538
0.56526     0.05725
0.53270     0.05870
0.50000     0.05975
0.46730     0.06036
0.43474     0.06055
0.40245     0.06032
0.37059     0.05971
0.33928     0.05871
0.30866     0.05738
0.27886     0.05573
0.25000     0.05380
0.22221     0.05160
0.19562     0.04913
0.17033     0.04640
0.14645     0.04344
0.12408     0.04027
0.10332     0.03690
0.08427     0.03335
0.06699     0.02966
0.05156     0.02585
0.03806     0.02196
0.02653     0.01807
0.01704     0.01427
0.00961     0.01060
0.00428     0.00706
0.00107     0.00354
0.00000     0.00000
0.00107     -0.00354
0.00428     -0.00706
0.00961     -0.01060
0.01704     -0.01427
0.02653     -0.01807
0.03806     -0.02196
0.05156     -0.02585
0.06699     -0.02966
0.08427     -0.03335
0.10332     -0.03690
0.12408     -0.04027
0.14645     -0.04344
0.17033     -0.04640
0.19562     -0.04913
0.22221     -0.05160
0.25000     -0.05380
0.27886     -0.05573
0.30866     -0.05738
0.33928     -0.05871
0.37059     -0.05971
0.40245     -0.06032
0.43474     -0.06055
0.46730     -0.06036
0.50000     -0.05975
0.53270     -0.05870
0.56526     -0.05725
0.59755     -0.05538
0.62941     -0.05308
0.66072     -0.05032
0.69134     -0.04707
0.72114     -0.04330
0.75000     -0.03903
0.77779     -0.03434
0.80438     -0.02941
0.82967     -0.02455
0.85355     -0.02003
0.87592     -0.01602
0.89668     -0.01260
0.91573     -0.00976
0.93301     -0.00746
0.94844     -0.00564
0.96194     -0.00422
0.97347     -0.00314
0.98296     -0.00229
0.99039     -0.00167
0.99572     -0.00112
0.99893     -0.00080
1.00000     0.00000
No parser warnings Send to airfoil plotter
Add to comparison
Lednicer format dat file
Selig format dat file

Similar airfoils

LWK 80-120/K25PreviewDetails
NASA/LANGLEY 64-012 AIRFOILPreviewDetails
NACA 64-012A AIRFOILPreviewDetails
NACA 0012-34PreviewDetails
NACA 0012-64PreviewDetails
EPPLER E836 HYDROFOIL AIRFOILPreviewDetails
S1012PreviewDetails
NASA SC(2)-0012 AIRFOILPreviewDetails
NACA 16-012PreviewDetails
NASA/LANGLEY LS(1)-0013 AIRFOILPreviewDetails

Polars for AH 85-L-120/17 (ah85l120-il)

PlotAirfoilReynolds #NcritMax Cl/CdDescriptionSource 
   ah85l120-il50,000921.4 at α=5.25°Mach=0 Ncrit=9Xfoil predictionDetails
   ah85l120-il50,000527.1 at α=4.5°Mach=0 Ncrit=5Xfoil predictionDetails
   ah85l120-il100,000937.6 at α=4°Mach=0 Ncrit=9Xfoil predictionDetails
   ah85l120-il100,000529.8 at α=4.5°Mach=0 Ncrit=5Xfoil predictionDetails
   ah85l120-il200,000535 at α=4°Mach=0 Ncrit=5Xfoil predictionDetails
   ah85l120-il500,000948.3 at α=3.75°Mach=0 Ncrit=9Xfoil predictionDetails
   ah85l120-il500,000543.5 at α=3°Mach=0 Ncrit=5Xfoil predictionDetails
   ah85l120-il1,000,000955.4 at α=3.25°Mach=0 Ncrit=9Xfoil predictionDetails
   ah85l120-il1,000,000552 at α=8.25°Mach=0 Ncrit=5Xfoil predictionDetails
Reynolds number calculator
Set Reynolds number and Ncrit rangeLowHigh
Reynolds Number
NCrit