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

WORTMANN FX 66-17AII-182 AIRFOIL (AS TESTED AT NASA) (fx6617ai-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: WORTMANN FX 66-17AII-182 AIRFOIL (AS TESTED AT NASA) (fx6617ai-il)
Reynolds number: 200,000
Max Cl/Cd: 68.84 at α=9.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-fx6617ai-il-200000-n5.txt
Download as CSV file: xf-fx6617ai-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: WORTMANN FX 66-17AII-182 AIRFOIL (AS TESTED AT N
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -13.250  -0.5574   0.07502   0.07030  -0.0671   0.7171   0.0175
 -13.000  -0.5856   0.06727   0.06226  -0.0723   0.7092   0.0174
 -12.750  -0.6044   0.06164   0.05637  -0.0758   0.7002   0.0174
 -12.500  -0.6180   0.05704   0.05153  -0.0783   0.6910   0.0175
 -12.250  -0.6278   0.05313   0.04740  -0.0801   0.6825   0.0176
 -12.000  -0.6337   0.04981   0.04386  -0.0815   0.6743   0.0177
 -11.750  -0.6366   0.04690   0.04075  -0.0825   0.6669   0.0179
 -11.500  -0.6368   0.04432   0.03796  -0.0832   0.6599   0.0180
 -11.250  -0.6345   0.04199   0.03542  -0.0838   0.6538   0.0183
 -11.000  -0.6294   0.03996   0.03319  -0.0842   0.6477   0.0186
 -10.750  -0.6233   0.03826   0.03129  -0.0842   0.6425   0.0190
 -10.500  -0.6152   0.03674   0.02956  -0.0838   0.6379   0.0195
 -10.250  -0.6035   0.03512   0.02772  -0.0835   0.6331   0.0201
 -10.000  -0.5897   0.03352   0.02585  -0.0831   0.6287   0.0206
  -9.750  -0.5737   0.03201   0.02404  -0.0827   0.6248   0.0211
  -9.500  -0.5552   0.03058   0.02238  -0.0823   0.6205   0.0215
  -9.250  -0.5355   0.02942   0.02096  -0.0820   0.6162   0.0219
  -9.000  -0.5154   0.02807   0.01957  -0.0817   0.6123   0.0226
  -8.750  -0.4944   0.02716   0.01857  -0.0815   0.6090   0.0233
  -8.500  -0.4726   0.02626   0.01759  -0.0813   0.6056   0.0240
  -8.250  -0.4504   0.02533   0.01658  -0.0810   0.6021   0.0245
  -8.000  -0.4283   0.02445   0.01561  -0.0807   0.5990   0.0251
  -7.750  -0.4061   0.02363   0.01467  -0.0804   0.5961   0.0258
  -7.500  -0.3839   0.02289   0.01380  -0.0800   0.5935   0.0265
  -7.250  -0.3612   0.02218   0.01300  -0.0798   0.5906   0.0273
  -7.000  -0.3392   0.02139   0.01217  -0.0796   0.5874   0.0281
  -6.750  -0.3170   0.02064   0.01139  -0.0794   0.5844   0.0292
  -6.500  -0.2935   0.02005   0.01073  -0.0794   0.5816   0.0307
  -6.250  -0.2692   0.01955   0.01013  -0.0793   0.5791   0.0328
  -6.000  -0.2446   0.01908   0.00953  -0.0793   0.5770   0.0351
  -5.750  -0.2202   0.01848   0.00891  -0.0794   0.5746   0.0383
  -5.500  -0.1946   0.01804   0.00840  -0.0795   0.5720   0.0421
  -5.250  -0.1694   0.01753   0.00788  -0.0796   0.5696   0.0487
  -5.000  -0.1440   0.01703   0.00742  -0.0797   0.5672   0.0648
  -4.750  -0.1199   0.01631   0.00691  -0.0799   0.5649   0.1148
  -4.500  -0.0968   0.01541   0.00639  -0.0802   0.5628   0.2098
  -4.250  -0.0742   0.01448   0.00597  -0.0803   0.5608   0.3329
  -4.000  -0.0494   0.01406   0.00596  -0.0802   0.5584   0.4362
  -3.750  -0.0224   0.01403   0.00604  -0.0802   0.5561   0.4864
  -3.500   0.0054   0.01411   0.00614  -0.0802   0.5539   0.5173
  -3.250   0.0333   0.01428   0.00630  -0.0801   0.5518   0.5420
  -3.000   0.0615   0.01447   0.00642  -0.0801   0.5499   0.5614
  -2.750   0.0899   0.01459   0.00645  -0.0802   0.5481   0.5712
  -2.500   0.1188   0.01465   0.00637  -0.0806   0.5463   0.5782
  -2.250   0.1473   0.01473   0.00639  -0.0808   0.5443   0.5846
  -2.000   0.1756   0.01482   0.00645  -0.0810   0.5420   0.5928
  -1.750   0.2040   0.01490   0.00647  -0.0812   0.5398   0.5996
  -1.500   0.2325   0.01496   0.00650  -0.0814   0.5379   0.6042
  -1.250   0.2612   0.01501   0.00647  -0.0818   0.5360   0.6091
  -1.000   0.2901   0.01505   0.00642  -0.0822   0.5343   0.6138
  -0.750   0.3186   0.01511   0.00644  -0.0825   0.5327   0.6173
  -0.500   0.3473   0.01518   0.00645  -0.0828   0.5312   0.6215
  -0.250   0.3761   0.01526   0.00646  -0.0832   0.5295   0.6259
   0.000   0.4044   0.01533   0.00652  -0.0835   0.5274   0.6301
   0.250   0.4324   0.01541   0.00663  -0.0838   0.5253   0.6336
   0.500   0.4606   0.01550   0.00672  -0.0840   0.5234   0.6377
   0.750   0.4891   0.01560   0.00678  -0.0844   0.5217   0.6420
   1.000   0.5177   0.01569   0.00684  -0.0848   0.5200   0.6460
   1.250   0.5459   0.01578   0.00694  -0.0851   0.5185   0.6494
   1.500   0.5743   0.01587   0.00701  -0.0854   0.5170   0.6534
   1.750   0.6030   0.01598   0.00707  -0.0858   0.5155   0.6577
   2.000   0.6313   0.01611   0.00720  -0.0861   0.5138   0.6617
   2.250   0.6584   0.01624   0.00742  -0.0863   0.5118   0.6649
   2.500   0.6859   0.01640   0.00763  -0.0865   0.5099   0.6689
   2.750   0.7136   0.01655   0.00782  -0.0868   0.5081   0.6733
   3.000   0.7416   0.01671   0.00799  -0.0871   0.5064   0.6775
   3.250   0.7692   0.01683   0.00816  -0.0874   0.5047   0.6807
   3.500   0.7971   0.01695   0.00831  -0.0876   0.5031   0.6847
   3.750   0.8253   0.01708   0.00845  -0.0880   0.5016   0.6892
   4.000   0.8538   0.01723   0.00859  -0.0884   0.5003   0.6935
   4.250   0.8809   0.01742   0.00886  -0.0885   0.4989   0.6970
   4.500   0.9068   0.01767   0.00923  -0.0886   0.4969   0.7013
   4.750   0.9331   0.01791   0.00957  -0.0887   0.4948   0.7060
   5.000   0.9595   0.01812   0.00986  -0.0888   0.4927   0.7103
   5.250   0.9860   0.01830   0.01014  -0.0889   0.4908   0.7145
   5.500   1.0131   0.01849   0.01040  -0.0891   0.4892   0.7194
   5.750   1.0405   0.01869   0.01064  -0.0894   0.4877   0.7247
   6.000   1.0678   0.01884   0.01088  -0.0896   0.4863   0.7291
   6.250   1.0958   0.01900   0.01109  -0.0899   0.4850   0.7343
   6.500   1.1203   0.01934   0.01156  -0.0898   0.4828   0.7401
   6.750   1.1436   0.01970   0.01211  -0.0895   0.4805   0.7450
   7.000   1.1678   0.02002   0.01257  -0.0894   0.4783   0.7509
   7.250   1.1927   0.02029   0.01296  -0.0893   0.4761   0.7569
   7.500   1.2181   0.02049   0.01328  -0.0892   0.4742   0.7630
   7.750   1.2445   0.02069   0.01358  -0.0894   0.4725   0.7703
   8.000   1.2713   0.02083   0.01382  -0.0894   0.4710   0.7772
   8.250   1.2944   0.02106   0.01419  -0.0891   0.4677   0.7853
   8.500   1.3137   0.02124   0.01460  -0.0881   0.4621   0.7944
   8.750   1.3423   0.02064   0.01399  -0.0880   0.4557   0.8054
   9.000   1.3582   0.02080   0.01439  -0.0865   0.4476   0.8194
   9.250   1.3818   0.02053   0.01422  -0.0857   0.4413   0.8383
   9.500   1.3963   0.02075   0.01473  -0.0839   0.4338   0.8793
   9.750   1.4187   0.02061   0.01471  -0.0833   0.4262   1.0002
  10.000   1.4316   0.02114   0.01539  -0.0816   0.4167   1.0002
  10.250   1.4420   0.02155   0.01587  -0.0794   0.4051   1.0002
  10.500   1.4491   0.02221   0.01657  -0.0771   0.3896   1.0002
  10.750   1.4511   0.02323   0.01754  -0.0745   0.3680   1.0002
  11.000   1.4430   0.02519   0.01939  -0.0718   0.3370   1.0002
  11.250   1.4221   0.02871   0.02273  -0.0693   0.3021   1.0002
  11.500   1.3945   0.03349   0.02735  -0.0675   0.2715   1.0002
  11.750   1.3650   0.03879   0.03253  -0.0661   0.2430   1.0002
  12.000   1.3328   0.04443   0.03803  -0.0646   0.2131   1.0002
  12.250   1.2998   0.05021   0.04364  -0.0633   0.1829   1.0002
  12.500   1.2715   0.05576   0.04903  -0.0623   0.1546   1.0002
  12.750   1.2494   0.06088   0.05400  -0.0616   0.1269   1.0002
  13.000   1.2293   0.06593   0.05888  -0.0611   0.0981   1.0002
  13.250   1.2145   0.07051   0.06332  -0.0608   0.0759   1.0002
  13.500   1.2062   0.07444   0.06717  -0.0606   0.0614   1.0002
  13.750   1.2018   0.07798   0.07070  -0.0605   0.0526   1.0002
  14.000   1.1988   0.08137   0.07409  -0.0604   0.0464   1.0002
  14.250   1.1994   0.08441   0.07718  -0.0604   0.0420   1.0002
  14.500   1.1988   0.08762   0.08041  -0.0605   0.0383   1.0002
  14.750   1.1990   0.09076   0.08361  -0.0606   0.0356   1.0002
  15.000   1.2011   0.09367   0.08659  -0.0607   0.0330   1.0002
  15.250   1.2012   0.09683   0.08980  -0.0610   0.0308   1.0002
<< Back to WORTMANN FX 66-17AII-182 AIRFOIL (AS TESTED AT NASA) (fx6617ai-il)

Polar data table (+)

Polar graphs


<< Back to WORTMANN FX 66-17AII-182 AIRFOIL (AS TESTED AT NASA) (fx6617ai-il)