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

DSMA-523B AIRFOIL (dsma523b-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: DSMA-523B AIRFOIL (dsma523b-il)
Reynolds number: 500,000
Max Cl/Cd: 49.56 at α=6.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-dsma523b-il-500000.txt
Download as CSV file: xf-dsma523b-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: DSMA-523B AIRFOIL                               
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -14.250  -0.7695   0.08154   0.07853  -0.0458   1.0000   0.0304
 -14.000  -0.7825   0.10797   0.10511  -0.0254   1.0000   0.0307
 -13.750  -0.9417   0.06939   0.06600  -0.0514   1.0000   0.0295
 -13.500  -0.9905   0.05995   0.05628  -0.0557   1.0000   0.0293
 -13.250  -1.0285   0.05327   0.04932  -0.0568   1.0000   0.0292
 -13.000  -1.0568   0.04853   0.04433  -0.0560   1.0000   0.0292
 -12.750  -1.0819   0.04505   0.04066  -0.0535   1.0000   0.0293
 -12.500  -1.0988   0.04179   0.03719  -0.0515   1.0000   0.0294
 -12.250  -1.1010   0.03834   0.03347  -0.0518   1.0000   0.0296
 -12.000  -1.0938   0.03538   0.03025  -0.0526   1.0000   0.0299
 -11.750  -1.0799   0.03296   0.02758  -0.0537   1.0000   0.0302
 -11.500  -1.0619   0.03111   0.02549  -0.0546   1.0000   0.0305
 -11.250  -1.0443   0.02897   0.02315  -0.0550   1.0000   0.0309
 -11.000  -1.0245   0.02744   0.02156  -0.0550   1.0000   0.0313
 -10.750  -1.0020   0.02647   0.02054  -0.0554   1.0000   0.0317
 -10.500  -0.9782   0.02564   0.01966  -0.0558   1.0000   0.0322
 -10.250  -0.9538   0.02481   0.01877  -0.0562   1.0000   0.0327
 -10.000  -0.9287   0.02399   0.01788  -0.0567   1.0000   0.0333
  -9.750  -0.9032   0.02315   0.01695  -0.0571   1.0000   0.0340
  -9.500  -0.8771   0.02235   0.01604  -0.0576   1.0000   0.0347
  -9.250  -0.8506   0.02157   0.01517  -0.0581   1.0000   0.0353
  -9.000  -0.8237   0.02078   0.01438  -0.0587   1.0000   0.0362
  -8.750  -0.7958   0.02038   0.01399  -0.0594   1.0000   0.0370
  -8.500  -0.7677   0.01999   0.01357  -0.0600   1.0000   0.0379
  -8.250  -0.7394   0.01957   0.01310  -0.0607   1.0000   0.0391
  -8.000  -0.7108   0.01911   0.01257  -0.0613   1.0000   0.0403
  -7.750  -0.6821   0.01877   0.01226  -0.0621   1.0000   0.0414
  -7.500  -0.6534   0.01859   0.01208  -0.0627   1.0000   0.0426
  -7.250  -0.6245   0.01836   0.01180  -0.0633   1.0000   0.0443
  -7.000  -0.5953   0.01804   0.01145  -0.0640   1.0000   0.0460
  -6.750  -0.5664   0.01783   0.01126  -0.0646   1.0000   0.0474
  -6.500  -0.5372   0.01762   0.01103  -0.0652   1.0000   0.0493
  -6.250  -0.5080   0.01748   0.01081  -0.0658   1.0000   0.0511
  -6.000  -0.4771   0.01694   0.01038  -0.0669   1.0000   0.0532
  -5.750  -0.4472   0.01672   0.01016  -0.0677   1.0000   0.0556
  -5.500  -0.4150   0.01633   0.00978  -0.0690   1.0000   0.0585
  -5.250  -0.3819   0.01596   0.00946  -0.0706   1.0000   0.0616
  -5.000  -0.3471   0.01557   0.00909  -0.0725   1.0000   0.0648
  -4.750  -0.3099   0.01514   0.00872  -0.0751   1.0000   0.0685
  -4.500  -0.2750   0.01489   0.00848  -0.0771   1.0000   0.0718
  -4.250  -0.2394   0.01465   0.00830  -0.0791   1.0000   0.0755
  -4.000  -0.2066   0.01455   0.00822  -0.0805   1.0000   0.0786
  -3.750  -0.1736   0.01445   0.00820  -0.0819   1.0000   0.0842
  -3.250  -0.1062   0.01428   0.00828  -0.0850   1.0000   0.1218
  -3.000  -0.0675   0.01408   0.00833  -0.0878   1.0000   0.1725
  -2.750  -0.0253   0.01383   0.00840  -0.0914   1.0000   0.2438
  -2.500   0.0269   0.01338   0.00842  -0.0975   1.0000   0.3627
  -2.250   0.0874   0.01285   0.00879  -0.1052   1.0000   0.5823
  -2.000   0.1145   0.01309   0.00910  -0.1048   0.9993   0.6132
  -1.750   0.1726   0.01290   0.00889  -0.1108   0.9940   0.6315
  -1.500   0.2229   0.01259   0.00861  -0.1150   0.9873   0.6415
  -1.250   0.2822   0.01198   0.00799  -0.1211   0.9799   0.6503
  -1.000   0.3328   0.01148   0.00757  -0.1250   0.9729   0.6574
  -0.750   0.3837   0.01104   0.00717  -0.1289   0.9662   0.6689
  -0.500   0.3676   0.01122   0.00744  -0.1185   0.9336   0.6720
  -0.250   0.4715   0.00990   0.00610  -0.1332   0.9073   0.6819
   0.000   0.4941   0.01232   0.00634  -0.1317   0.4239   0.6861
   0.250   0.5130   0.01331   0.00668  -0.1297   0.2667   0.6883
   0.500   0.5334   0.01422   0.00705  -0.1279   0.1441   0.6905
   0.750   0.5557   0.01481   0.00744  -0.1263   0.1030   0.6933
   1.000   0.5802   0.01522   0.00779  -0.1251   0.0905   0.6966
   1.250   0.6075   0.01562   0.00811  -0.1248   0.0815   0.7002
   1.500   0.6389   0.01594   0.00838  -0.1256   0.0763   0.7044
   1.750   0.6644   0.01624   0.00864  -0.1248   0.0711   0.7066
   2.000   0.6889   0.01651   0.00895  -0.1237   0.0684   0.7085
   2.250   0.7138   0.01676   0.00922  -0.1227   0.0654   0.7102
   2.500   0.7390   0.01707   0.00951  -0.1218   0.0625   0.7121
   2.750   0.7635   0.01763   0.01006  -0.1209   0.0598   0.7143
   3.000   0.7902   0.01790   0.01036  -0.1203   0.0580   0.7177
   3.250   0.8195   0.01816   0.01060  -0.1206   0.0557   0.7203
   3.500   0.8493   0.01844   0.01084  -0.1210   0.0537   0.7217
   3.750   0.8769   0.01922   0.01157  -0.1211   0.0515   0.7230
   4.000   0.9066   0.01949   0.01185  -0.1214   0.0504   0.7236
   4.250   0.9354   0.01977   0.01216  -0.1215   0.0489   0.7243
   4.500   0.9632   0.02004   0.01244  -0.1215   0.0475   0.7249
   4.750   0.9906   0.02034   0.01273  -0.1213   0.0461   0.7256
   5.000   1.0161   0.02125   0.01360  -0.1209   0.0444   0.7261
   5.250   1.0433   0.02161   0.01404  -0.1206   0.0434   0.7267
   5.500   1.0703   0.02195   0.01443  -0.1204   0.0422   0.7273
   5.750   1.0972   0.02228   0.01479  -0.1201   0.0409   0.7279
   6.000   1.1236   0.02268   0.01520  -0.1197   0.0399   0.7286
   6.250   1.1494   0.02319   0.01571  -0.1193   0.0390   0.7293
   6.500   1.1737   0.02465   0.01720  -0.1187   0.0378   0.7302
   6.750   1.1994   0.02513   0.01779  -0.1181   0.0373   0.7310
   7.000   1.2250   0.02578   0.01856  -0.1175   0.0366   0.7317
   7.250   1.2502   0.02652   0.01940  -0.1170   0.0359   0.7323
   7.500   1.2750   0.02727   0.02022  -0.1164   0.0352   0.7329
   7.750   1.2993   0.02803   0.02107  -0.1157   0.0346   0.7336
   8.000   1.3233   0.02875   0.02186  -0.1150   0.0340   0.7342
   8.250   1.3468   0.02947   0.02262  -0.1142   0.0335   0.7349
   8.500   1.3694   0.03047   0.02367  -0.1134   0.0330   0.7356
   8.750   1.3876   0.03308   0.02646  -0.1120   0.0323   0.7362
   9.000   1.4075   0.03408   0.02765  -0.1106   0.0320   0.7369
   9.250   1.4254   0.03545   0.02925  -0.1088   0.0317   0.7375
   9.500   1.4411   0.03709   0.03115  -0.1068   0.0312   0.7382
   9.750   1.4538   0.03907   0.03341  -0.1044   0.0308   0.7388
  10.000   1.4632   0.04131   0.03592  -0.1015   0.0304   0.7395
  10.250   1.4686   0.04385   0.03874  -0.0982   0.0300   0.7402
  10.500   1.4710   0.04639   0.04156  -0.0946   0.0297   0.7410
  10.750   1.4662   0.04944   0.04490  -0.0901   0.0295   0.7415
  11.000   1.4597   0.05192   0.04759  -0.0854   0.0292   0.7419
  11.250   1.4232   0.05633   0.05237  -0.0771   0.0292   0.7421
  11.500   1.3822   0.06182   0.05823  -0.0702   0.0292   0.7422
<< Back to DSMA-523B AIRFOIL (dsma523b-il)

Polar data table (+)

Polar graphs


<< Back to DSMA-523B AIRFOIL (dsma523b-il)