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

RAF 89 AIRFOIL (raf89-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: RAF 89 AIRFOIL (raf89-il)
Reynolds number: 100,000
Max Cl/Cd: 18.71 at α=9.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-raf89-il-100000.txt
Download as CSV file: xf-raf89-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAF 89 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -14.500  -0.3206   0.13664   0.13041  -0.0522   1.0000   0.1877
 -14.250  -0.8773   0.05861   0.05158  -0.0749   1.0000   0.2018
 -14.000  -0.9623   0.05497   0.04781  -0.0674   1.0000   0.2019
 -13.750  -1.0285   0.05252   0.04522  -0.0604   1.0000   0.2024
 -13.500  -1.0870   0.04988   0.04234  -0.0543   0.9991   0.2036
 -13.250  -0.9936   0.05271   0.04566  -0.0590   0.9965   0.2122
 -13.000  -1.0290   0.04818   0.04072  -0.0594   0.9870   0.2172
 -12.750  -0.9685   0.04914   0.04195  -0.0629   0.9828   0.2248
 -12.500  -0.9754   0.04671   0.03929  -0.0637   0.9759   0.2312
 -12.250  -0.9833   0.04497   0.03740  -0.0620   0.9665   0.2363
 -12.000  -0.9316   0.04545   0.03806  -0.0656   0.9622   0.2444
 -11.750  -0.9627   0.04350   0.03580  -0.0611   0.9519   0.2490
 -11.500  -0.9156   0.04366   0.03616  -0.0641   0.9473   0.2565
 -11.250  -0.9080   0.04284   0.03523  -0.0633   0.9416   0.2634
 -11.000  -0.9507   0.04209   0.03425  -0.0542   0.9299   0.2662
 -10.750  -0.8960   0.04211   0.03447  -0.0586   0.9262   0.2744
 -10.500  -0.8882   0.04148   0.03369  -0.0570   0.9208   0.2814
 -10.250  -0.9317   0.04110   0.03314  -0.0468   0.9099   0.2838
 -10.000  -0.8941   0.04104   0.03323  -0.0487   0.9055   0.2911
  -9.750  -0.8724   0.04065   0.03278  -0.0486   0.9009   0.2988
  -9.500  -0.9045   0.04021   0.03212  -0.0399   0.8919   0.3025
  -9.250  -0.8870   0.04014   0.03216  -0.0387   0.8863   0.3083
  -9.000  -0.8623   0.04008   0.03214  -0.0385   0.8816   0.3156
  -8.750  -0.8609   0.03917   0.03088  -0.0354   0.8769   0.3231
  -8.500  -0.8567   0.03941   0.03131  -0.0317   0.8697   0.3274
  -8.250  -0.8492   0.03951   0.03147  -0.0288   0.8636   0.3328
  -8.000  -0.8442   0.03899   0.03075  -0.0257   0.8582   0.3400
  -7.750  -0.8003   0.03858   0.03041  -0.0287   0.8548   0.3483
  -7.500  -0.8040   0.03875   0.03059  -0.0240   0.8484   0.3533
  -7.250  -0.8193   0.03863   0.03036  -0.0175   0.8407   0.3581
  -7.000  -0.8039   0.03811   0.02972  -0.0159   0.8353   0.3652
  -6.750  -0.7550   0.03793   0.02967  -0.0194   0.8318   0.3742
  -6.500  -0.7651   0.03789   0.02949  -0.0137   0.8250   0.3798
  -6.250  -0.7754   0.03774   0.02923  -0.0081   0.8169   0.3848
  -6.000  -0.7363   0.03763   0.02929  -0.0100   0.8118   0.3930
  -5.750  -0.7080   0.03710   0.02856  -0.0102   0.8077   0.4028
  -5.500  -0.7081   0.03728   0.02882  -0.0061   0.7998   0.4079
  -5.250  -0.6984   0.03739   0.02901  -0.0035   0.7922   0.4145
  -5.000  -0.6772   0.03695   0.02838  -0.0025   0.7867   0.4241
  -4.750  -0.6191   0.03670   0.02835  -0.0071   0.7834   0.4343
  -4.500  -0.6450   0.03722   0.02883   0.0008   0.7727   0.4389
  -4.250  -0.6282   0.03702   0.02858   0.0024   0.7657   0.4476
  -4.000  -0.5815   0.03689   0.02857  -0.0002   0.7614   0.4582
  -3.750  -0.5673   0.03694   0.02856   0.0019   0.7549   0.4675
  -3.500  -0.5704   0.03745   0.02918   0.0063   0.7443   0.4732
  -3.000  -0.4824   0.03713   0.02894   0.0023   0.7360   0.4967
  -2.750  -0.5165   0.03799   0.02974   0.0110   0.7220   0.5025
  -2.500  -0.4787   0.03798   0.02982   0.0100   0.7165   0.5131
  -2.250  -0.4365   0.03766   0.02942   0.0085   0.7132   0.5275
  -2.000  -0.4579   0.03877   0.03063   0.0151   0.6994   0.5325
  -1.750  -0.4291   0.03863   0.03040   0.0156   0.6938   0.5467
  -1.500  -0.3764   0.03843   0.03034   0.0128   0.6907   0.5590
  -1.250  -0.4076   0.03946   0.03128   0.0206   0.6763   0.5684
  -1.000  -0.3674   0.03939   0.03134   0.0195   0.6715   0.5803
  -0.750  -0.3247   0.03890   0.03082   0.0184   0.6685   0.5961
  -0.500  -0.3476   0.04020   0.03219   0.0248   0.6539   0.6041
  -0.250  -0.3139   0.03986   0.03184   0.0247   0.6494   0.6196
   0.000  -0.2713   0.03939   0.03138   0.0238   0.6467   0.6361
   0.250  -0.2965   0.04085   0.03291   0.0301   0.6318   0.6443
   0.500  -0.2613   0.04046   0.03251   0.0302   0.6278   0.6607
   0.750  -0.2103   0.03995   0.03211   0.0282   0.6255   0.6743
   1.000  -0.2486   0.04165   0.03375   0.0360   0.6102   0.6855
   1.250  -0.2034   0.04133   0.03355   0.0347   0.6068   0.6980
   1.500  -0.1594   0.04056   0.03276   0.0339   0.6047   0.7130
   1.750  -0.1915   0.04268   0.03495   0.0404   0.5893   0.7215
   2.000  -0.1570   0.04216   0.03440   0.0406   0.5861   0.7356
   2.250  -0.1071   0.04140   0.03369   0.0391   0.5843   0.7483
   2.500  -0.1503   0.04372   0.03599   0.0466   0.5686   0.7582
   2.750  -0.1050   0.04311   0.03544   0.0454   0.5659   0.7695
   3.000  -0.0815   0.04293   0.03521   0.0467   0.5618   0.7819
   3.250  -0.1007   0.04496   0.03733   0.0509   0.5484   0.7908
   3.500  -0.0661   0.04413   0.03645   0.0515   0.5458   0.8040
   3.750  -0.0250   0.04384   0.03622   0.0502   0.5423   0.8142
   4.000  -0.0594   0.04640   0.03882   0.0557   0.5283   0.8255
   4.250  -0.0078   0.04553   0.03798   0.0536   0.5261   0.8365
   4.500  -0.0396   0.04867   0.04118   0.0576   0.5121   0.8483
   4.750   0.0101   0.04829   0.04084   0.0549   0.5086   0.8598
   5.000   0.0606   0.04730   0.03985   0.0530   0.5066   0.8730
   5.250   0.0516   0.05126   0.04394   0.0521   0.4922   0.8837
   5.500   0.1020   0.05082   0.04351   0.0493   0.4890   0.8973
   5.750   0.1794   0.04979   0.04252   0.0433   0.4873   0.9055
   6.000   0.2297   0.04969   0.04244   0.0400   0.4837   0.9177
   6.250   0.3259   0.04713   0.03986   0.0331   0.4844   0.9245
   6.500   0.4160   0.04472   0.03744   0.0268   0.4844   0.9312
   6.750   0.4980   0.04200   0.03466   0.0219   0.4845   0.9390
   7.000   0.6061   0.03859   0.03114   0.0138   0.4844   0.9402
   7.250   0.5300   0.04731   0.04010   0.0170   0.4642   0.9572
   7.500   0.6236   0.04412   0.03686   0.0107   0.4638   0.9609
   7.750   0.7393   0.03991   0.03252   0.0024   0.4636   0.9625
   8.000   0.5137   0.06459   0.05767   0.0068   0.4242   0.9903
   8.250   0.6060   0.06064   0.05369   0.0011   0.4249   0.9982
   8.500   0.6223   0.05645   0.04935   0.0051   0.4259   1.0000
   8.750   0.6629   0.05261   0.04536   0.0071   0.4263   1.0000
   9.000   0.7267   0.04875   0.04138   0.0064   0.4264   1.0000
   9.250   0.8234   0.04400   0.03650   0.0022   0.4263   1.0000
  11.000   0.4102   0.09955   0.09220   0.0201   0.3494   1.0000
  11.250   0.3632   0.10774   0.10046   0.0191   0.3447   1.0000
  11.500   0.3587   0.11128   0.10401   0.0189   0.3411   1.0000
<< Back to RAF 89 AIRFOIL (raf89-il)

Polar data table (+)

Polar graphs


<< Back to RAF 89 AIRFOIL (raf89-il)