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

FX 66-182 AIRFOIL (fx66182-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: FX 66-182 AIRFOIL (fx66182-il)
Reynolds number: 50,000
Max Cl/Cd: 4.66 at α=10.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fx66182-il-50000.txt
Download as CSV file: xf-fx66182-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 66-182 AIRFOIL                               
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.500  -0.2510   0.13224   0.12657  -0.0437   1.0000   0.2705
 -11.250  -0.2895   0.13624   0.13101  -0.0365   1.0000   0.2703
 -11.000  -0.2999   0.13600   0.13088  -0.0348   0.9954   0.2737
 -10.750  -0.2840   0.13196   0.12682  -0.0402   0.9859   0.2881
 -10.500  -0.2700   0.12845   0.12328  -0.0449   0.9762   0.3023
 -10.250  -0.2296   0.12204   0.11677  -0.0493   0.9678   0.3153
 -10.000  -0.2055   0.11738   0.11206  -0.0538   0.9602   0.3302
  -9.750  -0.1850   0.11279   0.10742  -0.0578   0.9516   0.3396
  -9.500  -0.1816   0.10960   0.10423  -0.0627   0.9455   0.3501
  -9.250  -0.2866   0.08772   0.08248  -0.0864   0.9356   0.1573
  -9.000  -0.3419   0.07867   0.07344  -0.0953   0.9252   0.1440
  -8.750  -0.3771   0.07332   0.06797  -0.0987   0.9155   0.1376
  -8.500  -0.4369   0.07126   0.06556  -0.0958   0.9032   0.1315
  -8.250  -0.4332   0.06704   0.06115  -0.0970   0.8975   0.1294
  -8.000  -0.4510   0.06540   0.05938  -0.0935   0.8891   0.1277
  -7.750  -0.4519   0.06147   0.05503  -0.0939   0.8838   0.1248
  -7.500  -0.4725   0.06026   0.05357  -0.0891   0.8770   0.1230
  -7.250  -0.4796   0.05779   0.05055  -0.0868   0.8721   0.1203
  -7.000  -0.4656   0.05492   0.04718  -0.0868   0.8683   0.1202
  -6.750  -0.4642   0.05366   0.04579  -0.0841   0.8641   0.1222
  -6.500  -0.4611   0.05238   0.04425  -0.0816   0.8599   0.1240
  -6.250  -0.4459   0.05061   0.04212  -0.0807   0.8561   0.1258
  -6.000  -0.4201   0.04871   0.03978  -0.0809   0.8529   0.1281
  -5.750  -0.4052   0.04779   0.03842  -0.0796   0.8503   0.1320
  -5.500  -0.3984   0.04701   0.03767  -0.0771   0.8480   0.1362
  -5.250  -0.3855   0.04642   0.03703  -0.0752   0.8453   0.1417
  -5.000  -0.3681   0.04592   0.03629  -0.0736   0.8428   0.1474
  -4.750  -0.3510   0.04536   0.03590  -0.0719   0.8410   0.1562
  -4.500  -0.3378   0.04500   0.03562  -0.0700   0.8399   0.1695
  -4.250  -0.3254   0.04452   0.03525  -0.0681   0.8388   0.1900
  -4.000  -0.5726   0.04506   0.03557  -0.0330   1.0000   0.1339
  -3.750  -0.5540   0.04416   0.03475  -0.0320   1.0000   0.1402
  -3.500  -0.5345   0.04358   0.03400  -0.0308   1.0000   0.1468
  -3.250  -0.5170   0.04299   0.03357  -0.0293   1.0000   0.1555
  -3.000  -0.4995   0.04247   0.03312  -0.0281   1.0000   0.1710
  -2.750  -0.4812   0.04172   0.03258  -0.0274   1.0000   0.1984
  -2.500  -0.4717   0.03999   0.03367  -0.0235   1.0000   0.5498
  -2.250  -0.4752   0.04248   0.03611  -0.0138   1.0000   0.6545
  -2.000  -0.4764   0.04379   0.03738  -0.0059   1.0000   0.7030
  -1.750  -0.4786   0.04454   0.03811   0.0019   1.0000   0.7424
  -1.500  -0.4819   0.04495   0.03846   0.0097   1.0000   0.7819
  -1.250  -0.4879   0.04504   0.03852   0.0182   1.0000   0.8212
  -1.000  -0.4927   0.04498   0.03842   0.0264   1.0000   0.8636
  -0.750  -0.4840   0.04542   0.03877   0.0319   1.0000   0.9086
  -0.500  -0.4082   0.04878   0.04172   0.0227   1.0000   0.9447
  -0.250  -0.3601   0.05040   0.04301   0.0159   1.0000   0.9568
   0.000  -0.3023   0.05260   0.04488   0.0068   0.9957   0.9653
   0.250  -0.2503   0.05460   0.04660  -0.0013   0.9877   0.9730
   0.500  -0.2021   0.05649   0.04824  -0.0087   0.9785   0.9808
   0.750  -0.1544   0.05855   0.05008  -0.0161   0.9701   0.9893
   1.000  -0.1079   0.06057   0.05190  -0.0232   0.9600   0.9991
   1.250  -0.0891   0.06114   0.05233  -0.0247   0.9465   1.0000
   1.500  -0.0766   0.06137   0.05245  -0.0249   0.9327   1.0000
   1.750  -0.0689   0.06135   0.05234  -0.0241   0.9196   1.0000
   2.000  -0.0616   0.06147   0.05235  -0.0232   0.9076   1.0000
   2.250  -0.0400   0.06312   0.05384  -0.0250   0.8987   1.0000
   2.500  -0.0290   0.06296   0.05357  -0.0247   0.8852   1.0000
   2.750  -0.0138   0.06351   0.05402  -0.0252   0.8738   1.0000
   3.000   0.0229   0.06634   0.05668  -0.0296   0.8674   1.0000
   3.250   0.0359   0.06644   0.05668  -0.0297   0.8553   1.0000
   3.500   0.0582   0.06793   0.05807  -0.0316   0.8472   1.0000
   3.750   0.0897   0.07000   0.06002  -0.0348   0.8377   1.0000
   4.000   0.1057   0.07093   0.06087  -0.0355   0.8275   1.0000
   4.250   0.1417   0.07378   0.06360  -0.0394   0.8210   1.0000
   4.500   0.1537   0.07438   0.06414  -0.0394   0.8102   1.0000
   4.750   0.1929   0.07786   0.06752  -0.0437   0.8045   1.0000
   5.000   0.2011   0.07806   0.06767  -0.0431   0.7929   1.0000
   5.250   0.2346   0.08130   0.07083  -0.0465   0.7878   1.0000
   5.500   0.2458   0.08183   0.07132  -0.0463   0.7763   1.0000
   5.750   0.2660   0.08389   0.07333  -0.0476   0.7695   1.0000
   6.000   0.2883   0.08571   0.07512  -0.0490   0.7598   1.0000
   6.250   0.3035   0.08745   0.07682  -0.0495   0.7526   1.0000
   6.500   0.3282   0.08965   0.07900  -0.0512   0.7436   1.0000
   6.750   0.3398   0.09119   0.08052  -0.0513   0.7358   1.0000
   7.000   0.3654   0.09365   0.08296  -0.0530   0.7276   1.0000
   7.250   0.3748   0.09511   0.08442  -0.0529   0.7194   1.0000
   7.500   0.4013   0.09780   0.08712  -0.0546   0.7114   1.0000
   7.750   0.4082   0.09917   0.08850  -0.0543   0.7033   1.0000
   8.000   0.4350   0.10204   0.09138  -0.0560   0.6954   1.0000
   8.250   0.4400   0.10333   0.09270  -0.0555   0.6868   1.0000
   8.500   0.4668   0.10637   0.09578  -0.0572   0.6795   1.0000
   8.750   0.4703   0.10764   0.09708  -0.0567   0.6705   1.0000
   9.000   0.4981   0.11092   0.10041  -0.0585   0.6634   1.0000
   9.250   0.4993   0.11207   0.10160  -0.0578   0.6543   1.0000
   9.500   0.5283   0.11567   0.10528  -0.0597   0.6474   1.0000
   9.750   0.5272   0.11666   0.10632  -0.0589   0.6378   1.0000
  10.000   0.5570   0.12056   0.11030  -0.0608   0.6314   1.0000
  10.250   0.5541   0.12139   0.11119  -0.0600   0.6213   1.0000
  10.500   0.5860   0.12582   0.11572  -0.0621   0.6153   1.0000
<< Back to FX 66-182 AIRFOIL (fx66182-il)

Polar data table (+)

Polar graphs


<< Back to FX 66-182 AIRFOIL (fx66182-il)