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

RG 12A-1.8/9.0 AIRFOIL (rg12a189-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: RG 12A-1.8/9.0 AIRFOIL (rg12a189-il)
Reynolds number: 50,000
Max Cl/Cd: 35.9 at α=5.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-rg12a189-il-50000-n5.txt
Download as CSV file: xf-rg12a189-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RG 12A-1.8/9.0 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4940   0.09085   0.08384  -0.0331   1.0000   0.0408
  -8.750  -0.4950   0.08678   0.07984  -0.0342   1.0000   0.0403
  -8.500  -0.4983   0.08274   0.07586  -0.0357   1.0000   0.0398
  -8.250  -0.5057   0.07869   0.07188  -0.0371   1.0000   0.0393
  -8.000  -0.5121   0.07447   0.06770  -0.0383   1.0000   0.0389
  -7.750  -0.5179   0.07027   0.06347  -0.0391   1.0000   0.0384
  -7.500  -0.5220   0.06618   0.05932  -0.0396   1.0000   0.0379
  -7.250  -0.5248   0.06195   0.05498  -0.0398   1.0000   0.0376
  -7.000  -0.5249   0.05776   0.05061  -0.0396   1.0000   0.0372
  -6.750  -0.5219   0.05367   0.04627  -0.0392   1.0000   0.0369
  -6.500  -0.5159   0.04965   0.04188  -0.0386   1.0000   0.0368
  -6.250  -0.5066   0.04582   0.03767  -0.0378   1.0000   0.0368
  -6.000  -0.4942   0.04225   0.03366  -0.0368   1.0000   0.0371
  -5.750  -0.4790   0.03897   0.02986  -0.0357   1.0000   0.0386
  -5.500  -0.4607   0.03608   0.02630  -0.0345   1.0000   0.0408
  -5.250  -0.4423   0.03342   0.02339  -0.0336   1.0000   0.0430
  -5.000  -0.4219   0.03124   0.02094  -0.0325   1.0000   0.0452
  -4.750  -0.3999   0.02918   0.01857  -0.0312   1.0000   0.0478
  -4.500  -0.3769   0.02751   0.01648  -0.0298   1.0000   0.0533
  -4.250  -0.3559   0.02601   0.01495  -0.0287   1.0000   0.0609
  -4.000  -0.3337   0.02452   0.01329  -0.0271   1.0000   0.0690
  -3.750  -0.3126   0.02335   0.01211  -0.0259   1.0000   0.0865
  -3.500  -0.2909   0.02211   0.01092  -0.0249   1.0000   0.1184
  -3.250  -0.2706   0.02075   0.00993  -0.0240   1.0000   0.1794
  -3.000  -0.2516   0.01916   0.00912  -0.0233   1.0000   0.3093
  -2.750  -0.2416   0.01765   0.00908  -0.0194   1.0000   0.5692
  -2.500  -0.2354   0.01756   0.00939  -0.0129   1.0000   0.7676
  -2.250  -0.2229   0.01760   0.00938  -0.0082   1.0000   0.8587
  -2.000  -0.1686   0.01770   0.00915  -0.0114   1.0000   0.9504
  -1.750  -0.0934   0.01769   0.00859  -0.0207   1.0000   1.0000
  -1.500  -0.0900   0.01759   0.00834  -0.0174   1.0000   1.0000
  -1.250  -0.0833   0.01755   0.00814  -0.0147   1.0000   1.0000
  -1.000  -0.0722   0.01759   0.00800  -0.0129   1.0000   1.0000
  -0.750  -0.0440   0.01781   0.00802  -0.0141   0.9947   1.0000
  -0.500  -0.0002   0.01813   0.00809  -0.0182   0.9830   1.0000
  -0.250   0.0424   0.01839   0.00817  -0.0218   0.9706   1.0000
   0.000   0.0836   0.01860   0.00824  -0.0251   0.9574   1.0000
   0.250   0.1240   0.01877   0.00828  -0.0282   0.9437   1.0000
   0.500   0.1636   0.01889   0.00833  -0.0309   0.9295   1.0000
   0.750   0.2025   0.01899   0.00838  -0.0333   0.9149   1.0000
   1.000   0.2410   0.01905   0.00841  -0.0356   0.8998   1.0000
   1.250   0.2796   0.01908   0.00843  -0.0378   0.8845   1.0000
   1.500   0.3180   0.01908   0.00846  -0.0397   0.8688   1.0000
   1.750   0.3546   0.01907   0.00849  -0.0412   0.8521   1.0000
   2.000   0.3913   0.01903   0.00850  -0.0426   0.8351   1.0000
   2.250   0.4282   0.01897   0.00849  -0.0439   0.8177   1.0000
   2.500   0.4585   0.01900   0.00860  -0.0440   0.7965   1.0000
   2.750   0.4926   0.01897   0.00862  -0.0445   0.7765   1.0000
   3.000   0.5220   0.01902   0.00874  -0.0443   0.7535   1.0000
   3.250   0.5529   0.01907   0.00882  -0.0443   0.7307   1.0000
   3.500   0.5819   0.01917   0.00902  -0.0439   0.7060   1.0000
   3.750   0.6090   0.01934   0.00923  -0.0432   0.6797   1.0000
   4.000   0.6355   0.01954   0.00948  -0.0424   0.6525   1.0000
   4.250   0.6611   0.01978   0.00977  -0.0415   0.6241   1.0000
   4.500   0.6859   0.02007   0.01013  -0.0404   0.5945   1.0000
   4.750   0.7100   0.02039   0.01048  -0.0393   0.5638   1.0000
   5.000   0.7328   0.02076   0.01089  -0.0380   0.5312   1.0000
   5.250   0.7548   0.02118   0.01134  -0.0366   0.4970   1.0000
   5.500   0.7763   0.02165   0.01185  -0.0352   0.4617   1.0000
   5.750   0.7966   0.02219   0.01240  -0.0336   0.4242   1.0000
   6.000   0.8161   0.02281   0.01302  -0.0321   0.3853   1.0000
   6.250   0.8347   0.02352   0.01370  -0.0305   0.3449   1.0000
   6.500   0.8521   0.02434   0.01447  -0.0288   0.3030   1.0000
   6.750   0.8685   0.02532   0.01537  -0.0271   0.2601   1.0000
   7.000   0.8838   0.02647   0.01647  -0.0255   0.2174   1.0000
   7.250   0.8980   0.02784   0.01771  -0.0239   0.1765   1.0000
   7.500   0.9108   0.02944   0.01912  -0.0223   0.1395   1.0000
   7.750   0.9234   0.03118   0.02075  -0.0207   0.1085   1.0000
   8.000   0.9355   0.03305   0.02251  -0.0191   0.0886   1.0000
   8.250   0.9490   0.03495   0.02448  -0.0175   0.0740   1.0000
   8.500   0.9639   0.03699   0.02661  -0.0160   0.0660   1.0000
   8.750   0.9823   0.03926   0.02919  -0.0147   0.0598   1.0000
   9.000   0.9972   0.04131   0.03135  -0.0135   0.0544   1.0000
   9.250   1.0127   0.04378   0.03407  -0.0124   0.0500   1.0000
   9.500   1.0281   0.04664   0.03734  -0.0112   0.0472   1.0000
   9.750   1.0398   0.04967   0.04073  -0.0098   0.0453   1.0000
  10.000   1.0473   0.05281   0.04418  -0.0084   0.0439   1.0000
  10.250   1.0510   0.05608   0.04773  -0.0068   0.0428   1.0000
  10.500   1.0499   0.05942   0.05130  -0.0051   0.0419   1.0000
  11.000   1.0314   0.06683   0.05919  -0.0020   0.0408   1.0000
  11.250   1.0150   0.07099   0.06363  -0.0016   0.0407   1.0000
  11.500   0.9968   0.07573   0.06861  -0.0024   0.0406   1.0000
  11.750   0.9778   0.08114   0.07422  -0.0044   0.0407   1.0000
  12.000   0.9592   0.08723   0.08047  -0.0076   0.0409   1.0000
  12.250   0.9103   0.10036   0.09393  -0.0178   0.0435   1.0000
<< Back to RG 12A-1.8/9.0 AIRFOIL (rg12a189-il)

Polar data table (+)

Polar graphs


<< Back to RG 12A-1.8/9.0 AIRFOIL (rg12a189-il)