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-S-161 AIRFOIL (fx66s161-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: FX 66-S-161 AIRFOIL (fx66s161-il)
Reynolds number: 50,000
Max Cl/Cd: 5.32 at α=9°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fx66s161-il-50000.txt
Download as CSV file: xf-fx66s161-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 66-S-161 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
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.2431   0.11589   0.11002  -0.0367   1.0000   0.2865
  -9.750  -0.2420   0.11325   0.10753  -0.0363   1.0000   0.2946
  -9.500  -0.2583   0.11317   0.10769  -0.0351   1.0000   0.3044
  -9.250  -0.2967   0.11645   0.11132  -0.0271   1.0000   0.3038
  -9.000  -0.3388   0.11987   0.11496  -0.0209   0.9999   0.3044
  -8.750  -0.3122   0.11488   0.10993  -0.0264   0.9906   0.3196
  -8.500  -0.2773   0.10914   0.10410  -0.0315   0.9822   0.3316
  -8.250  -0.2640   0.07782   0.07309  -0.0720   0.9373   0.1583
  -8.000  -0.2559   0.07196   0.06719  -0.0751   0.9302   0.1507
  -7.750  -0.4065   0.07508   0.07008  -0.0760   0.9430   0.1397
  -7.500  -0.4454   0.06741   0.06192  -0.0825   0.9299   0.1286
  -7.250  -0.4326   0.06308   0.05741  -0.0842   0.9214   0.1246
  -7.000  -0.4240   0.05798   0.05187  -0.0871   0.9136   0.1201
  -6.750  -0.4243   0.05405   0.04728  -0.0875   0.9054   0.1162
  -6.500  -0.4038   0.05037   0.04298  -0.0893   0.8987   0.1160
  -6.250  -0.3874   0.04821   0.04058  -0.0892   0.8917   0.1196
  -6.000  -0.3658   0.04584   0.03773  -0.0897   0.8851   0.1226
  -5.750  -0.3359   0.04341   0.03467  -0.0908   0.8796   0.1261
  -5.500  -0.3246   0.04216   0.03328  -0.0891   0.8735   0.1311
  -5.250  -0.2941   0.04068   0.03148  -0.0898   0.8685   0.1415
  -5.000  -0.2713   0.03943   0.03028  -0.0892   0.8635   0.1525
  -4.750  -0.2581   0.03872   0.02957  -0.0874   0.8581   0.1665
  -4.500  -0.2312   0.03745   0.02846  -0.0872   0.8535   0.1928
  -4.250  -0.2116   0.03606   0.02757  -0.0867   0.8494   0.2483
  -4.000  -0.2137   0.03584   0.02893  -0.0818   0.8455   0.4263
  -3.750  -0.2143   0.03849   0.03173  -0.0740   0.8414   0.5370
  -3.500  -0.2031   0.04030   0.03343  -0.0687   0.8376   0.5909
  -3.250  -0.1992   0.04160   0.03469  -0.0629   0.8345   0.6242
  -3.000  -0.2010   0.04250   0.03554  -0.0575   0.8320   0.6527
  -2.750  -0.2006   0.04327   0.03626  -0.0522   0.8298   0.6818
  -2.500  -0.2010   0.04392   0.03689  -0.0464   0.8284   0.7088
  -2.250  -0.2011   0.04439   0.03732  -0.0412   0.8278   0.7354
  -2.000  -0.1983   0.04470   0.03753  -0.0372   0.8274   0.7605
  -1.750  -0.1944   0.04493   0.03767  -0.0336   0.8271   0.7823
  -1.500  -0.1886   0.04513   0.03774  -0.0313   0.8282   0.8024
  -1.250  -0.1814   0.04538   0.03786  -0.0294   0.8311   0.8197
  -1.000  -0.1692   0.04570   0.03804  -0.0284   0.8335   0.8361
  -0.750  -0.1726   0.04589   0.03819  -0.0262   0.8446   0.8501
  -0.500  -0.1514   0.04666   0.03880  -0.0271   0.8510   0.8663
   0.000  -0.2932   0.04176   0.03446  -0.0013   1.0000   0.8842
   0.250  -0.2695   0.04225   0.03482  -0.0028   1.0000   0.9061
   0.500  -0.2376   0.04300   0.03546  -0.0064   1.0000   0.9314
   0.750  -0.1989   0.04386   0.03619  -0.0122   0.9987   0.9627
   1.000  -0.1509   0.04573   0.03781  -0.0204   0.9871   1.0000
   1.250  -0.1100   0.04767   0.03947  -0.0271   0.9768   1.0000
   1.500  -0.0685   0.04981   0.04132  -0.0335   0.9665   1.0000
   1.750  -0.0225   0.05255   0.04377  -0.0406   0.9566   1.0000
   2.000   0.0120   0.05418   0.04517  -0.0452   0.9435   1.0000
   2.250   0.0445   0.05588   0.04666  -0.0491   0.9302   1.0000
   2.500   0.0751   0.05767   0.04824  -0.0525   0.9173   1.0000
   2.750   0.1048   0.05965   0.05002  -0.0556   0.9066   1.0000
   3.000   0.1442   0.06261   0.05277  -0.0601   0.8966   1.0000
   3.250   0.1672   0.06395   0.05396  -0.0616   0.8828   1.0000
   3.500   0.1885   0.06540   0.05527  -0.0627   0.8698   1.0000
   3.750   0.2097   0.06707   0.05682  -0.0638   0.8582   1.0000
   4.000   0.2419   0.06988   0.05950  -0.0666   0.8501   1.0000
   4.250   0.2651   0.07158   0.06111  -0.0678   0.8373   1.0000
   4.500   0.2799   0.07283   0.06229  -0.0677   0.8250   1.0000
   4.750   0.2977   0.07457   0.06396  -0.0682   0.8147   1.0000
   5.000   0.3326   0.07776   0.06706  -0.0711   0.8065   1.0000
   5.250   0.3432   0.07870   0.06798  -0.0704   0.7939   1.0000
   5.500   0.3564   0.08029   0.06953  -0.0703   0.7839   1.0000
   5.750   0.3899   0.08353   0.07273  -0.0729   0.7765   1.0000
   6.000   0.3975   0.08446   0.07365  -0.0719   0.7643   1.0000
   6.250   0.4111   0.08630   0.07549  -0.0719   0.7551   1.0000
   6.500   0.4398   0.08918   0.07835  -0.0737   0.7468   1.0000
   6.750   0.4461   0.09035   0.07954  -0.0728   0.7357   1.0000
   7.000   0.4730   0.09357   0.08276  -0.0745   0.7294   1.0000
   7.250   0.4838   0.09493   0.08416  -0.0742   0.7179   1.0000
   7.500   0.4926   0.09668   0.08594  -0.0738   0.7087   1.0000
   7.750   0.5225   0.10000   0.08929  -0.0757   0.7009   1.0000
   8.000   0.5243   0.10108   0.09041  -0.0746   0.6902   1.0000
   8.250   0.5545   0.10495   0.09434  -0.0766   0.6843   1.0000
   8.500   0.5580   0.10593   0.09539  -0.0756   0.6726   1.0000
   8.750   0.5658   0.10801   0.09751  -0.0755   0.6643   1.0000
   9.000   0.5921   0.11124   0.10082  -0.0769   0.6559   1.0000
   9.250   0.5930   0.11264   0.10228  -0.0761   0.6457   1.0000
   9.750   0.6228   0.11779   0.10762  -0.0772   0.6280   1.0000
  10.000   0.6314   0.12025   0.11016  -0.0773   0.6197   1.0000
  10.250   0.6559   0.12364   0.11366  -0.0785   0.6107   1.0000
  10.500   0.6552   0.12520   0.11530  -0.0781   0.6004   1.0000
  10.750   0.6816   0.12946   0.11970  -0.0796   0.5939   1.0000
  11.000   0.6845   0.13093   0.12127  -0.0793   0.5822   1.0000
  11.250   0.6881   0.13325   0.12368  -0.0795   0.5732   1.0000
  11.500   0.7186   0.13779   0.12838  -0.0811   0.5649   1.0000
<< Back to FX 66-S-161 AIRFOIL (fx66s161-il)

Polar data table (+)

Polar graphs


<< Back to FX 66-S-161 AIRFOIL (fx66s161-il)