Airfoil Tools
Search 1638 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

AH 93-K-132/15 (ah93k132-il)

AH 93-K-132/15 - Althaus AH 93-K-132/15 airfoil for use with flaps (K)


Airfoil ah93k132-il
Details Dat file Parser  
(ah93k132-il) AH 93-K-132/15
Althaus AH 93-K-132/15 airfoil for use with flaps (K)
Max thickness 13.2% at 41.2% chord.
Max camber 3.5% at 41.2% chord
Source UIUC Airfoil Coordinates Database
Source dat file
The dat file is in Selig format
AH 93-K-132/15
1.00000   0.00071
0.97897   0.00192
0.96313   0.00352
0.94521   0.00573
0.92646   0.00832
0.90745   0.01114
0.88859   0.01406
0.87068   0.01711
0.85388   0.02083
0.83732   0.02571
0.82000   0.03169
0.80183   0.03818
0.78351   0.04477
0.76535   0.05131
0.74750   0.05749
0.72983   0.06316
0.71216   0.06827
0.69434   0.07285
0.67628   0.07696
0.65800   0.08065
0.63952   0.08398
0.62089   0.08696
0.60214   0.08962
0.58331   0.09198
0.56440   0.09406
0.54543   0.09586
0.52641   0.09738
0.50734   0.09866
0.48823   0.09969
0.46911   0.10048
0.44998   0.10102
0.43087   0.10131
0.41175   0.10135
0.39264   0.10114
0.37355   0.10067
0.35448   0.09994
0.33543   0.09894
0.31640   0.09767
0.29740   0.09613
0.27845   0.09432
0.25955   0.09223
0.24073   0.08985
0.22200   0.08715
0.20337   0.08414
0.18485   0.08078
0.16647   0.07706
0.14825   0.07297
0.13024   0.06848
0.11249   0.06356
0.09507   0.05818
0.07809   0.05231
0.06178   0.04597
0.04656   0.03919
0.03307   0.03220
0.02212   0.02548
0.01404   0.01951
0.00841   0.01444
0.00455   0.01014
0.00199   0.00639
0.00049   0.00300
0.00000   0.00003
0.00068  -0.00288
0.00291  -0.00542
0.00645  -0.00763
0.01135  -0.00974
0.01820  -0.01194
0.02792  -0.01432
0.04134  -0.01680
0.05781  -0.01918
0.07581  -0.02125
0.09454  -0.02300
0.11356  -0.02444
0.13267  -0.02565
0.15185  -0.02664
0.17117  -0.02745
0.19062  -0.02812
0.21011  -0.02869
0.22964  -0.02915
0.24925  -0.02955
0.26886  -0.02989
0.28845  -0.03016
0.30810  -0.03038
0.32779  -0.03057
0.34746  -0.03073
0.36708  -0.03086
0.38671  -0.03091
0.40639  -0.03093
0.42611  -0.03093
0.44581  -0.03094
0.46552  -0.03096
0.48520  -0.03098
0.50483  -0.03095
0.52449  -0.03087
0.55644  -0.03070
0.57639  -0.03050
0.59632  -0.03025
0.61624  -0.02993
0.63616  -0.02954
0.65606  -0.02909
0.67595  -0.02856
0.69582  -0.02796
0.71566  -0.02726
0.73546  -0.02646
0.75521  -0.02555
0.77490  -0.02452
0.79452  -0.02337
0.81399  -0.02213
0.83331  -0.02078
0.85241  -0.01933
0.87006  -0.01788
0.88577  -0.01646
0.90023  -0.01502
0.91399  -0.01348
0.92753  -0.01177
0.94171  -0.00964
0.95644  -0.00692
0.96941  -0.00398
0.98074  -0.00187
0.99147  -0.00051
1.00000  -0.00014
No parser warnings Send to airfoil plotter
Add to comparison
Lednicer format dat file
Selig format dat file

Similar airfoils

AH 93-K-131/15PreviewDetails
AH 93-K-130/15PreviewDetails
AH 79-K-132/20PreviewDetails
AH 79-K-135/20 BPreviewDetails
74-130 WP2 MODPreviewDetails
AH 63-K-127/24PreviewDetails
AH 80-129PreviewDetails
AH 81-131PreviewDetails
AH 88-K-130/20PreviewDetails
74-130 WP2PreviewDetails

Polars for AH 93-K-132/15 (ah93k132-il)

PlotAirfoilReynolds #NcritMax Cl/CdDescriptionSource 
   ah93k132-il50,000925.8 at α=10.75°Mach=0 Ncrit=9Xfoil predictionDetails
   ah93k132-il50,000522.8 at α=10.25°Mach=0 Ncrit=5Xfoil predictionDetails
   ah93k132-il100,000942.5 at α=9.25°Mach=0 Ncrit=9Xfoil predictionDetails
   ah93k132-il100,000540 at α=8.5°Mach=0 Ncrit=5Xfoil predictionDetails
   ah93k132-il200,000973.4 at α=7.5°Mach=0 Ncrit=9Xfoil predictionDetails
   ah93k132-il200,000571.7 at α=6.5°Mach=0 Ncrit=5Xfoil predictionDetails
   ah93k132-il500,0009114.6 at α=6°Mach=0 Ncrit=9Xfoil predictionDetails
   ah93k132-il500,000599.1 at α=5°Mach=0 Ncrit=5Xfoil predictionDetails
   ah93k132-il1,000,0009134.6 at α=5.25°Mach=0 Ncrit=9Xfoil predictionDetails
   ah93k132-il1,000,0005103.4 at α=4.5°Mach=0 Ncrit=5Xfoil predictionDetails
Reynolds number calculator
Set Reynolds number and Ncrit rangeLowHigh
Reynolds Number
NCrit