Airfoil Tools
Search 1638 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

GIII BL430 AIRFOIL (giiim-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: GIII BL430 AIRFOIL (giiim-il)
Reynolds number: 100,000
Max Cl/Cd: 41.44 at α=4.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-giiim-il-100000-n5.txt
Download as CSV file: xf-giiim-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GIII BL430 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5397   0.10142   0.09621  -0.0007   1.0000   0.0507
  -8.500  -0.5426   0.09681   0.09164  -0.0067   1.0000   0.0508
  -8.250  -0.5452   0.09232   0.08711  -0.0121   1.0000   0.0509
  -8.000  -0.5445   0.08809   0.08278  -0.0159   1.0000   0.0510
  -7.750  -0.5312   0.08386   0.07870  -0.0136   1.0000   0.0516
  -7.500  -0.5222   0.08033   0.07519  -0.0134   1.0000   0.0522
  -7.250  -0.5144   0.07669   0.07155  -0.0146   1.0000   0.0530
  -7.000  -0.5061   0.07298   0.06780  -0.0162   1.0000   0.0539
  -6.750  -0.4968   0.06935   0.06410  -0.0179   1.0000   0.0554
  -6.500  -0.4869   0.06600   0.06033  -0.0219   1.0000   0.0593
  -6.250  -0.4761   0.06190   0.05607  -0.0226   1.0000   0.0601
  -6.000  -0.4627   0.05823   0.05255  -0.0222   1.0000   0.0615
  -5.750  -0.4483   0.05525   0.04953  -0.0221   1.0000   0.0631
  -5.500  -0.4329   0.05229   0.04644  -0.0221   1.0000   0.0647
  -5.250  -0.4164   0.04948   0.04343  -0.0220   1.0000   0.0667
  -5.000  -0.3970   0.04464   0.03807  -0.0211   1.0000   0.0498
  -4.750  -0.3810   0.04163   0.03493  -0.0202   1.0000   0.0472
  -4.500  -0.3638   0.03917   0.03226  -0.0191   1.0000   0.0476
  -4.250  -0.3465   0.03690   0.02975  -0.0177   1.0000   0.0480
  -4.000  -0.3291   0.03457   0.02718  -0.0162   1.0000   0.0470
  -3.750  -0.3111   0.03231   0.02461  -0.0145   1.0000   0.0460
  -3.500  -0.2930   0.03039   0.02242  -0.0129   1.0000   0.0457
  -3.250  -0.2599   0.02864   0.02043  -0.0143   0.9933   0.0477
  -3.000  -0.2220   0.02674   0.01817  -0.0162   0.9849   0.0489
  -2.750  -0.1848   0.02495   0.01606  -0.0179   0.9757   0.0491
  -2.500  -0.1472   0.02364   0.01433  -0.0193   0.9659   0.0508
  -2.250  -0.1120   0.02222   0.01292  -0.0210   0.9556   0.0528
  -2.000  -0.0754   0.02104   0.01161  -0.0225   0.9454   0.0537
  -1.750  -0.0406   0.02000   0.01050  -0.0236   0.9329   0.0548
  -1.500  -0.0069   0.01922   0.00963  -0.0243   0.9191   0.0576
  -1.250   0.0261   0.01855   0.00885  -0.0248   0.9049   0.0596
  -1.000   0.0569   0.01763   0.00801  -0.0251   0.8905   0.0609
  -0.750   0.0859   0.01695   0.00741  -0.0249   0.8748   0.0629
  -0.500   0.1132   0.01646   0.00696  -0.0244   0.8579   0.0666
  -0.250   0.1396   0.01607   0.00654  -0.0236   0.8402   0.0695
   0.000   0.1648   0.01562   0.00609  -0.0226   0.8223   0.0719
   0.250   0.1905   0.01531   0.00576  -0.0218   0.8036   0.0762
   0.500   0.2163   0.01511   0.00546  -0.0208   0.7822   0.0822
   0.750   0.2413   0.01485   0.00514  -0.0196   0.7557   0.0887
   1.000   0.2663   0.01465   0.00486  -0.0184   0.7262   0.1009
   1.250   0.2880   0.01331   0.00458  -0.0174   0.6987   0.4664
   1.750   0.3756   0.01223   0.00461  -0.0210   0.6461   1.0000
   2.000   0.3995   0.01238   0.00459  -0.0200   0.6218   1.0000
   2.250   0.4232   0.01256   0.00460  -0.0189   0.5974   1.0000
   2.500   0.4469   0.01276   0.00463  -0.0179   0.5714   1.0000
   2.750   0.4705   0.01299   0.00468  -0.0168   0.5456   1.0000
   3.000   0.4943   0.01322   0.00479  -0.0159   0.5187   1.0000
   3.250   0.5181   0.01348   0.00491  -0.0149   0.4914   1.0000
   3.500   0.5419   0.01375   0.00505  -0.0140   0.4616   1.0000
   3.750   0.5656   0.01403   0.00522  -0.0131   0.4277   1.0000
   4.000   0.5893   0.01434   0.00540  -0.0123   0.3861   1.0000
   4.250   0.6121   0.01477   0.00558  -0.0114   0.3330   1.0000
   4.500   0.6343   0.01535   0.00586  -0.0105   0.2863   1.0000
   4.750   0.6568   0.01598   0.00626  -0.0097   0.2502   1.0000
   5.000   0.6796   0.01659   0.00671  -0.0090   0.2221   1.0000
   5.250   0.7026   0.01720   0.00721  -0.0083   0.1996   1.0000
   5.500   0.7257   0.01780   0.00774  -0.0077   0.1797   1.0000
   5.750   0.7488   0.01840   0.00828  -0.0070   0.1631   1.0000
   6.000   0.7719   0.01902   0.00884  -0.0063   0.1487   1.0000
   6.250   0.7946   0.01968   0.00944  -0.0057   0.1369   1.0000
   6.500   0.8169   0.02039   0.01008  -0.0050   0.1267   1.0000
   6.750   0.8399   0.02104   0.01078  -0.0043   0.1172   1.0000
   7.000   0.8620   0.02181   0.01152  -0.0036   0.1096   1.0000
   7.250   0.8843   0.02255   0.01230  -0.0029   0.1024   1.0000
   7.500   0.9060   0.02342   0.01316  -0.0022   0.0966   1.0000
   7.750   0.9281   0.02422   0.01402  -0.0015   0.0907   1.0000
   8.000   0.9489   0.02520   0.01495  -0.0008   0.0864   1.0000
   8.250   0.9710   0.02613   0.01605  -0.0001   0.0818   1.0000
   8.500   0.9920   0.02706   0.01702   0.0006   0.0777   1.0000
   8.750   1.0122   0.02817   0.01812   0.0013   0.0744   1.0000
   9.000   1.0331   0.02933   0.01949   0.0021   0.0711   1.0000
   9.250   1.0532   0.03046   0.02073   0.0029   0.0681   1.0000
   9.500   1.0725   0.03158   0.02185   0.0036   0.0656   1.0000
   9.750   1.0911   0.03300   0.02343   0.0044   0.0632   1.0000
  10.000   1.1089   0.03452   0.02521   0.0052   0.0609   1.0000
  10.250   1.1260   0.03604   0.02690   0.0061   0.0590   1.0000
  10.500   1.1425   0.03743   0.02838   0.0069   0.0573   1.0000
  10.750   1.1587   0.03888   0.02983   0.0077   0.0558   1.0000
  11.000   1.1697   0.04098   0.03227   0.0088   0.0542   1.0000
  11.250   1.1780   0.04325   0.03489   0.0101   0.0527   1.0000
  11.500   1.1840   0.04563   0.03755   0.0113   0.0515   1.0000
  11.750   1.1872   0.04807   0.04024   0.0126   0.0506   1.0000
  12.000   1.1862   0.05047   0.04286   0.0141   0.0498   1.0000
  12.250   1.1835   0.05300   0.04557   0.0152   0.0492   1.0000
  12.500   1.1797   0.05570   0.04843   0.0158   0.0486   1.0000
  12.750   1.1760   0.05859   0.05144   0.0159   0.0480   1.0000
  13.000   1.1719   0.06176   0.05470   0.0156   0.0475   1.0000
  13.250   1.1514   0.06707   0.06028   0.0135   0.0472   1.0000
  13.500   1.1241   0.07406   0.06755   0.0095   0.0472   1.0000
  13.750   1.0917   0.08285   0.07659   0.0036   0.0473   1.0000
<< Back to GIII BL430 AIRFOIL (giiim-il)

Polar data table (+)

Polar graphs


<< Back to GIII BL430 AIRFOIL (giiim-il)