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RAE 5215 AIRFOIL (rae5215-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: RAE 5215 AIRFOIL (rae5215-il)
Reynolds number: 50,000
Max Cl/Cd: 26.66 at α=3°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-rae5215-il-50000-n5.txt
Download as CSV file: xf-rae5215-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAE 5215 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.5767   0.08989   0.08228  -0.0276   1.0000   0.0549
  -9.000  -0.5796   0.08485   0.07727  -0.0305   1.0000   0.0543
  -8.750  -0.5873   0.07964   0.07208  -0.0340   1.0000   0.0537
  -8.500  -0.5970   0.07436   0.06676  -0.0378   1.0000   0.0530
  -8.250  -0.6046   0.06920   0.06148  -0.0407   1.0000   0.0523
  -8.000  -0.6090   0.06422   0.05630  -0.0430   1.0000   0.0517
  -7.750  -0.6088   0.05954   0.05134  -0.0445   1.0000   0.0512
  -7.500  -0.6041   0.05515   0.04662  -0.0454   1.0000   0.0510
  -7.250  -0.5952   0.05113   0.04223  -0.0459   1.0000   0.0510
  -7.000  -0.5826   0.04766   0.03838  -0.0460   1.0000   0.0517
  -6.750  -0.5677   0.04442   0.03470  -0.0460   1.0000   0.0532
  -6.500  -0.5505   0.04127   0.03096  -0.0457   1.0000   0.0547
  -6.250  -0.5311   0.03865   0.02803  -0.0452   1.0000   0.0559
  -6.000  -0.5107   0.03650   0.02572  -0.0446   1.0000   0.0572
  -5.750  -0.4896   0.03461   0.02361  -0.0438   1.0000   0.0597
  -5.500  -0.4672   0.03273   0.02134  -0.0430   1.0000   0.0625
  -5.250  -0.4451   0.03099   0.01946  -0.0419   1.0000   0.0645
  -5.000  -0.4232   0.02959   0.01804  -0.0408   1.0000   0.0673
  -4.750  -0.4010   0.02838   0.01666  -0.0395   1.0000   0.0715
  -4.500  -0.3797   0.02720   0.01550  -0.0380   1.0000   0.0750
  -4.250  -0.3585   0.02619   0.01447  -0.0366   1.0000   0.0791
  -4.000  -0.3374   0.02525   0.01349  -0.0353   1.0000   0.0851
  -3.750  -0.3160   0.02435   0.01256  -0.0342   1.0000   0.0925
  -3.500  -0.2939   0.02340   0.01165  -0.0336   1.0000   0.1026
  -3.250  -0.2708   0.02241   0.01071  -0.0333   1.0000   0.1205
  -3.000  -0.2447   0.02088   0.00962  -0.0342   1.0000   0.1744
  -2.750  -0.2272   0.01912   0.00999  -0.0325   1.0000   0.5494
  -2.500  -0.2115   0.01960   0.01055  -0.0288   1.0000   0.6470
  -2.250  -0.2001   0.02004   0.01107  -0.0239   1.0000   0.7105
  -2.000  -0.1948   0.02030   0.01145  -0.0172   1.0000   0.7634
  -1.750  -0.1854   0.02025   0.01142  -0.0121   1.0000   0.7993
  -1.500  -0.1644   0.02008   0.01111  -0.0107   1.0000   0.8154
  -1.250  -0.1411   0.01992   0.01081  -0.0101   1.0000   0.8270
  -1.000  -0.1175   0.01977   0.01056  -0.0097   1.0000   0.8382
  -0.750  -0.0929   0.01967   0.01036  -0.0096   1.0000   0.8501
  -0.500  -0.0690   0.01958   0.01021  -0.0093   1.0000   0.8622
  -0.250  -0.0453   0.01951   0.01009  -0.0089   1.0000   0.8747
   0.000  -0.0213   0.01946   0.01003  -0.0088   1.0000   0.8885
   0.250   0.0032   0.01945   0.01002  -0.0088   1.0000   0.9041
   0.500   0.0289   0.01947   0.01006  -0.0091   1.0000   0.9230
   0.750   0.0561   0.01953   0.01016  -0.0101   1.0000   0.9505
   1.000   0.0972   0.01968   0.01035  -0.0140   0.9897   1.0000
   1.250   0.1423   0.02007   0.01073  -0.0186   0.9811   1.0000
   1.500   0.2004   0.02039   0.01108  -0.0250   0.9629   1.0000
   1.750   0.2981   0.01964   0.01045  -0.0355   0.9037   1.0000
   2.000   0.3503   0.01903   0.00991  -0.0383   0.8648   1.0000
   2.250   0.3873   0.01870   0.00966  -0.0390   0.8319   1.0000
   2.500   0.4206   0.01839   0.00942  -0.0388   0.7902   1.0000
   2.750   0.4495   0.01813   0.00918  -0.0377   0.7243   1.0000
   3.000   0.4820   0.01808   0.00826  -0.0355   0.5075   1.0000
   3.500   0.5191   0.02105   0.00933  -0.0329   0.2671   1.0000
   3.750   0.5439   0.02210   0.01007  -0.0329   0.2372   1.0000
   4.000   0.5702   0.02301   0.01082  -0.0329   0.2173   1.0000
   4.250   0.5972   0.02392   0.01159  -0.0330   0.2018   1.0000
   4.500   0.6253   0.02479   0.01243  -0.0332   0.1884   1.0000
   4.750   0.6542   0.02567   0.01332  -0.0334   0.1764   1.0000
   5.000   0.6831   0.02667   0.01422  -0.0336   0.1663   1.0000
   5.250   0.7118   0.02764   0.01522  -0.0337   0.1562   1.0000
   5.500   0.7405   0.02878   0.01640  -0.0339   0.1474   1.0000
   5.750   0.7682   0.02994   0.01756  -0.0339   0.1393   1.0000
   6.000   0.7957   0.03131   0.01905  -0.0339   0.1316   1.0000
   6.250   0.8221   0.03264   0.02046  -0.0339   0.1245   1.0000
   6.500   0.8476   0.03423   0.02220  -0.0337   0.1179   1.0000
   6.750   0.8722   0.03576   0.02392  -0.0334   0.1113   1.0000
   7.000   0.8960   0.03745   0.02570  -0.0332   0.1060   1.0000
   7.250   0.9180   0.03942   0.02807  -0.0326   0.1001   1.0000
   7.500   0.9404   0.04101   0.02967  -0.0322   0.0959   1.0000
   7.750   0.9584   0.04366   0.03283  -0.0314   0.0915   1.0000
   8.000   0.9759   0.04611   0.03562  -0.0305   0.0877   1.0000
   8.250   0.9949   0.04805   0.03765  -0.0300   0.0849   1.0000
   8.500   1.0057   0.05137   0.04143  -0.0288   0.0821   1.0000
   8.750   1.0111   0.05526   0.04586  -0.0275   0.0795   1.0000
   9.000   1.0160   0.05893   0.04991  -0.0263   0.0778   1.0000
   9.250   1.0200   0.06245   0.05371  -0.0253   0.0764   1.0000
   9.500   1.0280   0.06519   0.05657  -0.0244   0.0751   1.0000
   9.750   1.0264   0.06898   0.06057  -0.0234   0.0740   1.0000
  10.000   1.0037   0.07471   0.06668  -0.0225   0.0736   1.0000
  10.250   0.9764   0.08025   0.07245  -0.0220   0.0735   1.0000
  10.500   0.9475   0.08682   0.07917  -0.0239   0.0737   1.0000
  10.750   0.9188   0.09500   0.08746  -0.0285   0.0739   1.0000
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