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

Cambered plate C=8% T=5% R=1.6 (cp-080-050-gn) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: Cambered plate C=8% T=5% R=1.6 (cp-080-050-gn)
Reynolds number: 500,000
Max Cl/Cd: 44.82 at α=6.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-cp-080-050-gn-500000-n5.txt
Download as CSV file: xf-cp-080-050-gn-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: Cambered plate C=8% T=5% R=1.6                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.1297   0.10190   0.09886  -0.0946   0.9670   0.0340
  -8.000  -0.1180   0.09970   0.09667  -0.0958   0.9642   0.0342
  -7.750  -0.1022   0.09785   0.09483  -0.0974   0.9622   0.0346
  -7.500  -0.0878   0.09532   0.09231  -0.0996   0.9605   0.0349
  -7.250  -0.0727   0.09265   0.08964  -0.1021   0.9590   0.0353
  -7.000  -0.0574   0.08976   0.08674  -0.1050   0.9576   0.0358
  -6.750  -0.0412   0.08667   0.08365  -0.1083   0.9564   0.0359
  -6.500  -0.0348   0.08424   0.08122  -0.1090   0.9509   0.0364
  -6.250  -0.0272   0.07975   0.07671  -0.1133   0.9469   0.0374
  -6.000  -0.0069   0.07770   0.07466  -0.1159   0.9449   0.0376
  -5.750   0.0166   0.07565   0.07262  -0.1192   0.9434   0.0380
  -5.500   0.0401   0.07290   0.06987  -0.1237   0.9418   0.0384
  -5.250   0.0484   0.07090   0.06787  -0.1243   0.9349   0.0387
  -5.000   0.0646   0.06829   0.06525  -0.1274   0.9303   0.0391
  -4.750   0.0839   0.06529   0.06223  -0.1318   0.9264   0.0396
  -4.500   0.0808   0.06104   0.05793  -0.1354   0.9158   0.0409
  -4.250   0.1019   0.05920   0.05609  -0.1378   0.9118   0.0411
  -4.000   0.1202   0.05752   0.05440  -0.1393   0.9050   0.0414
  -3.750   0.1348   0.05585   0.05273  -0.1403   0.8981   0.0417
  -3.500   0.1564   0.05362   0.05044  -0.1432   0.8901   0.0420
  -3.250   0.1706   0.05142   0.04818  -0.1445   0.8761   0.0423
  -3.000   0.1857   0.04931   0.04603  -0.1457   0.8667   0.0428
  -2.750   0.2049   0.04688   0.04351  -0.1479   0.8563   0.0434
  -2.500   0.2197   0.04248   0.03897  -0.1516   0.8468   0.0448
  -2.250   0.2412   0.04116   0.03759  -0.1522   0.8381   0.0450
  -2.000   0.2603   0.03991   0.03632  -0.1521   0.8288   0.0452
  -1.750   0.2810   0.03842   0.03475  -0.1525   0.8153   0.0456
  -1.500   0.3015   0.03669   0.03292  -0.1529   0.8016   0.0459
  -1.000   0.3467   0.02927   0.02504  -0.1564   0.7793   0.0486
  -0.750   0.3652   0.02854   0.02419  -0.1550   0.7522   0.0487
  -0.500   0.3812   0.02793   0.02338  -0.1530   0.7164   0.0489
  -0.250   0.3921   0.02745   0.02262  -0.1498   0.6641   0.0492
   0.250   0.3631   0.02917   0.02254  -0.1335   0.2572   0.0494
   0.500   0.3654   0.02947   0.02219  -0.1290   0.0632   0.0497
   0.750   0.3856   0.02846   0.02108  -0.1278   0.0561   0.0503
   1.000   0.4064   0.02686   0.01934  -0.1267   0.0537   0.0511
   1.250   0.4281   0.02174   0.01362  -0.1258   0.0527   0.0531
   1.500   0.4480   0.02156   0.01341  -0.1242   0.0507   0.0533
   1.750   0.4678   0.02137   0.01318  -0.1224   0.0491   0.0536
   2.000   0.4883   0.02089   0.01263  -0.1208   0.0481   0.0541
   2.250   0.5090   0.02007   0.01166  -0.1192   0.0471   0.0547
   2.500   0.5300   0.01906   0.01042  -0.1176   0.0463   0.0554
   2.750   0.5512   0.01804   0.00913  -0.1160   0.0457   0.0564
   3.000   0.5723   0.01743   0.00825  -0.1143   0.0451   0.0574
   3.250   0.5924   0.01735   0.00813  -0.1125   0.0445   0.0577
   3.500   0.6119   0.01741   0.00816  -0.1106   0.0439   0.0580
   3.750   0.6324   0.01738   0.00810  -0.1089   0.0437   0.0583
   4.000   0.6532   0.01738   0.00806  -0.1072   0.0433   0.0586
   4.250   0.6735   0.01744   0.00809  -0.1054   0.0430   0.0590
   4.500   0.6935   0.01749   0.00811  -0.1036   0.0427   0.0595
   4.750   0.7133   0.01757   0.00815  -0.1017   0.0423   0.0598
   5.000   0.7327   0.01769   0.00824  -0.0998   0.0420   0.0603
   5.250   0.7519   0.01784   0.00835  -0.0978   0.0416   0.0609
   5.500   0.7707   0.01803   0.00851  -0.0958   0.0412   0.0615
   5.750   0.7897   0.01823   0.00868  -0.0938   0.0409   0.0619
   6.000   0.8085   0.01843   0.00886  -0.0918   0.0404   0.0622
   6.250   0.8267   0.01864   0.00906  -0.0898   0.0401   0.0624
   6.500   0.8443   0.01892   0.00936  -0.0877   0.0398   0.0629
   6.750   0.8618   0.01923   0.00969  -0.0856   0.0395   0.0632
   7.000   0.8787   0.01962   0.01008  -0.0834   0.0393   0.0635
   7.250   0.8952   0.02008   0.01054  -0.0812   0.0390   0.0639
   7.500   0.9118   0.02057   0.01104  -0.0790   0.0388   0.0643
   7.750   0.9301   0.02094   0.01144  -0.0771   0.0386   0.0650
   8.000   0.9483   0.02137   0.01190  -0.0752   0.0385   0.0657
   8.250   0.9669   0.02178   0.01233  -0.0734   0.0383   0.0662
   8.500   0.9855   0.02222   0.01281  -0.0716   0.0381   0.0667
   8.750   1.0043   0.02269   0.01330  -0.0698   0.0379   0.0672
   9.000   1.0233   0.02315   0.01380  -0.0681   0.0376   0.0677
   9.250   1.0424   0.02361   0.01430  -0.0664   0.0373   0.0684
   9.500   1.0617   0.02410   0.01483  -0.0648   0.0369   0.0691
   9.750   1.0812   0.02461   0.01538  -0.0633   0.0366   0.0701
  10.000   1.1011   0.02516   0.01596  -0.0618   0.0364   0.0712
  10.250   1.1210   0.02572   0.01655  -0.0603   0.0361   0.0725
  10.500   1.1412   0.02631   0.01717  -0.0589   0.0360   0.0736
  10.750   1.1606   0.02687   0.01778  -0.0575   0.0357   0.0760
  11.000   1.1804   0.02749   0.01842  -0.0560   0.0355   0.0797
  11.250   1.1999   0.02809   0.01907  -0.0546   0.0353   0.0864
  11.500   1.2192   0.02861   0.01986  -0.0532   0.0351   0.2427
  12.000   1.2922   0.02952   0.02194  -0.0577   0.0346   1.0000
  12.250   1.3144   0.03043   0.02283  -0.0567   0.0344   1.0000
  12.500   1.3369   0.03144   0.02387  -0.0558   0.0342   1.0000
  12.750   1.3540   0.03232   0.02484  -0.0541   0.0340   1.0000
  13.000   1.3697   0.03323   0.02585  -0.0523   0.0338   1.0000
  13.250   1.3846   0.03421   0.02693  -0.0504   0.0335   1.0000
  13.500   1.4001   0.03529   0.02812  -0.0487   0.0332   1.0000
  13.750   1.4157   0.03648   0.02940  -0.0470   0.0330   1.0000
  14.000   1.4298   0.03773   0.03076  -0.0452   0.0327   1.0000
  14.250   1.4423   0.03904   0.03219  -0.0432   0.0325   1.0000
  14.500   1.4530   0.04039   0.03365  -0.0412   0.0322   1.0000
  14.750   1.4628   0.04182   0.03519  -0.0391   0.0320   1.0000
  15.000   1.4709   0.04326   0.03673  -0.0369   0.0318   1.0000
  15.250   1.4777   0.04477   0.03836  -0.0347   0.0316   1.0000
  15.500   1.4835   0.04630   0.03998  -0.0324   0.0314   1.0000
  15.750   1.4881   0.04791   0.04169  -0.0302   0.0313   1.0000
  16.000   1.4911   0.04971   0.04361  -0.0280   0.0312   1.0000
  16.250   1.4931   0.05158   0.04559  -0.0259   0.0311   1.0000
  16.500   1.4935   0.05358   0.04770  -0.0239   0.0309   1.0000
  16.750   1.4932   0.05566   0.04989  -0.0219   0.0308   1.0000
  17.000   1.4927   0.05773   0.05207  -0.0202   0.0307   1.0000
  17.250   1.4905   0.06007   0.05452  -0.0185   0.0306   1.0000
  17.500   1.4889   0.06230   0.05683  -0.0171   0.0305   1.0000
  17.750   1.4838   0.06513   0.05978  -0.0158   0.0305   1.0000
  18.000   1.4813   0.06764   0.06237  -0.0148   0.0304   1.0000
  18.250   1.4744   0.07090   0.06575  -0.0140   0.0303   1.0000
  18.500   1.4672   0.07427   0.06924  -0.0136   0.0302   1.0000
  18.750   1.4577   0.07812   0.07321  -0.0134   0.0301   1.0000
  19.000   1.4443   0.08270   0.07794  -0.0138   0.0300   1.0000
  19.250   1.4082   0.09087   0.08644  -0.0161   0.0299   1.0000
<< Back to Cambered plate C=8% T=5% R=1.6 (cp-080-050-gn)

Polar data table (+)

Polar graphs


<< Back to Cambered plate C=8% T=5% R=1.6 (cp-080-050-gn)