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

NLR-7223-43 AIRFOIL (nl722343-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: NLR-7223-43 AIRFOIL (nl722343-il)
Reynolds number: 1,000,000
Max Cl/Cd: 82.56 at α=6.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-nl722343-il-1000000.txt
Download as CSV file: xf-nl722343-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NLR-7223-43 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.5077   0.08994   0.08825  -0.0244   1.0000   0.0099
  -9.250  -0.5091   0.08577   0.08409  -0.0264   1.0000   0.0101
  -9.000  -0.5138   0.08088   0.07923  -0.0290   1.0000   0.0102
  -7.250  -0.5048   0.02519   0.02291  -0.0489   0.9904   0.0111
  -7.000  -0.4845   0.02205   0.01963  -0.0508   0.9881   0.0113
  -6.750  -0.4615   0.01889   0.01630  -0.0526   0.9864   0.0115
  -6.500  -0.4367   0.01579   0.01300  -0.0544   0.9851   0.0119
  -6.250  -0.4171   0.01307   0.01004  -0.0544   0.9804   0.0126
  -6.000  -0.3958   0.00826   0.00450  -0.0544   0.9774   0.0144
  -5.750  -0.3673   0.00698   0.00316  -0.0556   0.9760   0.0148
  -5.500  -0.3377   0.00609   0.00218  -0.0568   0.9747   0.0159
  -5.250  -0.3389   0.01357   0.00859  -0.0537   0.9747   0.0114
  -5.000  -0.3067   0.01237   0.00727  -0.0547   0.9739   0.0114
  -4.750  -0.2787   0.01175   0.00662  -0.0549   0.9707   0.0118
  -4.500  -0.2510   0.01088   0.00569  -0.0549   0.9673   0.0119
  -4.250  -0.2211   0.01015   0.00492  -0.0554   0.9648   0.0121
  -4.000  -0.1899   0.00955   0.00429  -0.0562   0.9622   0.0126
  -3.750  -0.1587   0.00906   0.00376  -0.0569   0.9590   0.0132
  -3.500  -0.1321   0.00870   0.00337  -0.0566   0.9524   0.0137
  -3.250  -0.1026   0.00822   0.00284  -0.0569   0.9481   0.0144
  -3.000  -0.0747   0.00786   0.00246  -0.0569   0.9417   0.0154
  -2.750  -0.0459   0.00762   0.00217  -0.0569   0.9304   0.0168
  -2.500  -0.0200   0.00738   0.00186  -0.0562   0.9143   0.0193
  -2.250   0.0059   0.00718   0.00162  -0.0555   0.8992   0.0254
  -2.000   0.0304   0.00668   0.00140  -0.0549   0.8889   0.1126
  -1.750   0.0527   0.00576   0.00119  -0.0544   0.8813   0.3382
  -1.500   0.0731   0.00477   0.00114  -0.0533   0.8734   0.6499
  -1.250   0.1009   0.00487   0.00119  -0.0531   0.8655   0.6729
  -1.000   0.1280   0.00497   0.00129  -0.0526   0.8569   0.6909
  -0.750   0.1547   0.00510   0.00144  -0.0521   0.8474   0.7075
  -0.500   0.1813   0.00524   0.00157  -0.0515   0.8366   0.7216
  -0.250   0.2081   0.00534   0.00164  -0.0509   0.8233   0.7293
   0.000   0.2351   0.00544   0.00166  -0.0506   0.8073   0.7328
   0.250   0.2617   0.00543   0.00159  -0.0502   0.7871   0.7353
   0.500   0.2877   0.00550   0.00155  -0.0496   0.7566   0.7363
   0.750   0.3084   0.00586   0.00152  -0.0480   0.6570   0.7367
   1.000   0.3285   0.00640   0.00162  -0.0464   0.5465   0.7369
   1.250   0.3509   0.00686   0.00171  -0.0454   0.4546   0.7372
   1.500   0.3746   0.00725   0.00180  -0.0447   0.3794   0.7376
   1.750   0.3991   0.00759   0.00188  -0.0442   0.3175   0.7380
   2.000   0.4241   0.00790   0.00198  -0.0437   0.2655   0.7383
   2.250   0.4496   0.00817   0.00207  -0.0433   0.2263   0.7385
   2.500   0.4756   0.00839   0.00217  -0.0430   0.2002   0.7388
   2.750   0.5019   0.00858   0.00227  -0.0428   0.1828   0.7391
   3.000   0.5287   0.00872   0.00238  -0.0426   0.1731   0.7393
   3.250   0.5555   0.00886   0.00249  -0.0424   0.1662   0.7396
   3.500   0.5824   0.00898   0.00259  -0.0422   0.1611   0.7400
   3.750   0.6090   0.00915   0.00273  -0.0420   0.1551   0.7403
   4.000   0.6358   0.00927   0.00287  -0.0418   0.1518   0.7407
   4.250   0.6628   0.00937   0.00299  -0.0416   0.1484   0.7411
   4.500   0.6895   0.00951   0.00311  -0.0414   0.1436   0.7415
   4.750   0.7158   0.00969   0.00328  -0.0412   0.1379   0.7419
   5.000   0.7428   0.00977   0.00339  -0.0411   0.1337   0.7424
   5.250   0.7692   0.00994   0.00353  -0.0409   0.1278   0.7429
   5.500   0.7957   0.01008   0.00370  -0.0406   0.1229   0.7435
   5.750   0.8223   0.01021   0.00383  -0.0404   0.1166   0.7442
   6.000   0.8486   0.01038   0.00399  -0.0402   0.1087   0.7448
   6.250   0.8743   0.01059   0.00416  -0.0399   0.0938   0.7453
   6.500   0.8986   0.01099   0.00442  -0.0394   0.0729   0.7458
   6.750   0.9223   0.01146   0.00480  -0.0388   0.0551   0.7464
   7.000   0.9452   0.01204   0.00526  -0.0381   0.0347   0.7469
   7.250   0.9681   0.01265   0.00579  -0.0373   0.0232   0.7475
   7.500   0.9922   0.01308   0.00624  -0.0367   0.0202   0.7483
   7.750   1.0159   0.01356   0.00678  -0.0360   0.0180   0.7493
   8.000   1.0400   0.01394   0.00723  -0.0355   0.0167   0.7503
   8.250   1.0634   0.01444   0.00776  -0.0348   0.0154   0.7513
   8.500   1.0851   0.01514   0.00856  -0.0339   0.0142   0.7523
   8.750   1.1086   0.01557   0.00906  -0.0332   0.0137   0.7534
   9.000   1.1314   0.01609   0.00965  -0.0324   0.0131   0.7546
   9.250   1.1537   0.01663   0.01026  -0.0316   0.0126   0.7559
   9.500   1.1753   0.01723   0.01093  -0.0308   0.0121   0.7573
   9.750   1.1952   0.01799   0.01177  -0.0296   0.0117   0.7587
  10.000   1.2097   0.01931   0.01323  -0.0277   0.0111   0.7602
  10.250   1.2267   0.02029   0.01431  -0.0262   0.0110   0.7618
  10.500   1.2465   0.02092   0.01505  -0.0251   0.0108   0.7638
  10.750   1.2637   0.02176   0.01600  -0.0236   0.0106   0.7661
  11.000   1.2807   0.02257   0.01693  -0.0222   0.0104   0.7688
  11.250   1.2955   0.02353   0.01802  -0.0204   0.0102   0.7718
  11.500   1.3102   0.02443   0.01903  -0.0187   0.0099   0.7753
  11.750   1.3219   0.02551   0.02024  -0.0166   0.0098   0.7796
  12.000   1.3338   0.02626   0.02110  -0.0145   0.0095   0.7857
  12.250   1.3407   0.02723   0.02220  -0.0115   0.0093   0.7944
  12.500   1.3452   0.02829   0.02344  -0.0085   0.0092   0.8133
  12.750   1.3525   0.02922   0.02482  -0.0063   0.0091   1.0000
  13.000   1.3553   0.03058   0.02627  -0.0036   0.0089   1.0000
  13.250   1.3542   0.03230   0.02811  -0.0009   0.0089   1.0000
  13.500   1.3519   0.03420   0.03012   0.0015   0.0088   1.0000
  13.750   1.3458   0.03658   0.03262   0.0037   0.0087   1.0000
  14.000   1.3401   0.03911   0.03530   0.0053   0.0087   1.0000
  14.250   1.3279   0.04252   0.03884   0.0064   0.0086   1.0000
  14.500   1.3143   0.04649   0.04297   0.0067   0.0085   1.0000
  14.750   1.3013   0.05090   0.04753   0.0057   0.0085   1.0000
  15.000   1.2868   0.05619   0.05298   0.0035   0.0085   1.0000
  15.250   1.2721   0.06232   0.05926   0.0000   0.0085   1.0000
  15.500   1.2585   0.06896   0.06606  -0.0043   0.0085   1.0000
  15.750   1.2397   0.07702   0.07428  -0.0095   0.0085   1.0000
  16.000   1.2191   0.08566   0.08306  -0.0151   0.0085   1.0000
  16.250   1.2028   0.09371   0.09125  -0.0201   0.0086   1.0000
  16.500   1.1823   0.10264   0.10030  -0.0256   0.0087   1.0000
  16.750   1.1573   0.11253   0.11031  -0.0316   0.0087   1.0000
  17.000   1.1274   0.12421   0.12217  -0.0387   0.0088   1.0000
  17.250   1.0944   0.13712   0.13524  -0.0465   0.0090   1.0000
  17.500   1.0418   0.15633   0.15468  -0.0579   0.0092   1.0000
<< Back to NLR-7223-43 AIRFOIL (nl722343-il)

Polar data table (+)

Polar graphs


<< Back to NLR-7223-43 AIRFOIL (nl722343-il)