Cambered plate C=6% T=5% R=2.11 (cp-060-050-gn) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: Cambered plate C=6% T=5% R=2.11 (cp-060-050-gn) Reynolds number: 500,000 Max Cl/Cd: 45.25 at α=12.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-cp-060-050-gn-500000-n5.txt Download as CSV file: xf-cp-060-050-gn-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: Cambered plate C=6% T=5% R=2.11
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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.3153 0.11377 0.11078 -0.0473 0.9930 0.0429
-7.750 -0.3042 0.11088 0.10790 -0.0492 0.9915 0.0434
-7.250 -0.3036 0.10015 0.09713 -0.0562 0.9872 0.0455
-7.000 -0.2887 0.09789 0.09488 -0.0581 0.9855 0.0457
-6.750 -0.2724 0.09578 0.09278 -0.0600 0.9839 0.0459
-6.500 -0.2562 0.09337 0.09039 -0.0623 0.9824 0.0461
-6.250 -0.2338 0.09063 0.08766 -0.0661 0.9804 0.0466
-6.000 -0.2244 0.08805 0.08509 -0.0675 0.9755 0.0470
-5.750 -0.2114 0.08475 0.08179 -0.0706 0.9717 0.0476
-5.500 -0.2249 0.07631 0.07330 -0.0780 0.9689 0.0495
-5.250 -0.2046 0.07439 0.07139 -0.0802 0.9641 0.0496
-5.000 -0.1855 0.07234 0.06935 -0.0825 0.9566 0.0499
-4.750 -0.1633 0.06970 0.06672 -0.0865 0.9530 0.0502
-4.500 -0.1494 0.06714 0.06416 -0.0893 0.9450 0.0505
-4.000 -0.1463 0.05149 0.04829 -0.1103 0.9268 0.0536
-3.750 -0.1186 0.04948 0.04623 -0.1136 0.9201 0.0538
-3.500 -0.1010 0.04774 0.04445 -0.1145 0.9090 0.0540
-3.250 -0.0800 0.04551 0.04215 -0.1165 0.8987 0.0543
-2.750 -0.0780 0.03318 0.02920 -0.1190 0.8787 0.0578
-2.500 -0.0567 0.03204 0.02797 -0.1186 0.8660 0.0580
-1.500 0.0000 0.02034 0.01497 -0.1133 0.8011 0.0624
-1.250 0.0204 0.01999 0.01443 -0.1116 0.7630 0.0626
-1.000 0.0351 0.01987 0.01398 -0.1087 0.6943 0.0628
-0.750 0.0336 0.02049 0.01378 -0.1026 0.5138 0.0629
-0.500 0.0245 0.02218 0.01401 -0.0958 0.0883 0.0630
-0.250 0.0471 0.02185 0.01358 -0.0946 0.0768 0.0632
0.000 0.0698 0.02142 0.01306 -0.0935 0.0723 0.0636
0.250 0.0927 0.02090 0.01244 -0.0923 0.0698 0.0640
0.500 0.1155 0.02034 0.01178 -0.0911 0.0677 0.0646
0.750 0.1377 0.01951 0.01075 -0.0898 0.0663 0.0655
1.000 0.1592 0.01844 0.00940 -0.0883 0.0652 0.0665
1.250 0.1823 0.01779 0.00856 -0.0870 0.0644 0.0671
1.500 0.2061 0.01753 0.00827 -0.0858 0.0636 0.0674
1.750 0.2299 0.01737 0.00809 -0.0847 0.0628 0.0676
2.000 0.2537 0.01723 0.00793 -0.0834 0.0621 0.0679
2.250 0.2773 0.01716 0.00783 -0.0822 0.0614 0.0683
2.500 0.3009 0.01708 0.00773 -0.0809 0.0609 0.0687
2.750 0.3241 0.01702 0.00763 -0.0796 0.0603 0.0692
3.000 0.3469 0.01695 0.00751 -0.0781 0.0598 0.0697
3.250 0.3691 0.01689 0.00740 -0.0766 0.0593 0.0701
3.500 0.3918 0.01680 0.00726 -0.0751 0.0590 0.0706
3.750 0.4146 0.01673 0.00715 -0.0736 0.0587 0.0710
4.000 0.4372 0.01668 0.00706 -0.0721 0.0582 0.0715
4.250 0.4596 0.01668 0.00702 -0.0705 0.0577 0.0718
4.500 0.4818 0.01671 0.00701 -0.0689 0.0573 0.0722
4.750 0.5034 0.01668 0.00701 -0.0672 0.0569 0.0725
5.000 0.5248 0.01675 0.00710 -0.0655 0.0565 0.0730
5.250 0.5458 0.01685 0.00721 -0.0637 0.0562 0.0734
5.500 0.5662 0.01697 0.00733 -0.0617 0.0559 0.0739
5.750 0.5862 0.01708 0.00745 -0.0597 0.0556 0.0743
6.000 0.6062 0.01721 0.00757 -0.0577 0.0553 0.0747
6.250 0.6263 0.01736 0.00772 -0.0557 0.0550 0.0751
6.500 0.6463 0.01753 0.00789 -0.0537 0.0547 0.0755
6.750 0.6664 0.01775 0.00809 -0.0518 0.0545 0.0760
7.000 0.6866 0.01801 0.00834 -0.0499 0.0541 0.0764
7.250 0.7073 0.01835 0.00867 -0.0481 0.0538 0.0768
7.500 0.7287 0.01858 0.00892 -0.0464 0.0535 0.0774
7.750 0.7500 0.01881 0.00917 -0.0447 0.0533 0.0779
8.000 0.7715 0.01907 0.00948 -0.0430 0.0530 0.0785
8.250 0.7931 0.01936 0.00980 -0.0414 0.0528 0.0789
8.500 0.8149 0.01969 0.01016 -0.0399 0.0525 0.0795
8.750 0.8368 0.02002 0.01053 -0.0383 0.0522 0.0801
9.000 0.8592 0.02040 0.01094 -0.0369 0.0519 0.0807
9.250 0.8814 0.02078 0.01134 -0.0354 0.0516 0.0814
9.500 0.9036 0.02117 0.01176 -0.0339 0.0514 0.0821
9.750 0.9259 0.02158 0.01220 -0.0325 0.0511 0.0829
10.000 0.9478 0.02199 0.01264 -0.0310 0.0508 0.0837
10.250 0.9697 0.02241 0.01310 -0.0295 0.0505 0.0848
10.500 0.9908 0.02282 0.01354 -0.0279 0.0503 0.0864
10.750 1.0113 0.02322 0.01396 -0.0263 0.0500 0.0882
11.000 1.0322 0.02367 0.01443 -0.0247 0.0497 0.0900
11.250 1.0528 0.02413 0.01491 -0.0231 0.0495 0.0924
11.500 1.0737 0.02463 0.01543 -0.0215 0.0493 0.0962
11.750 1.0958 0.02522 0.01605 -0.0202 0.0490 0.1040
12.000 1.1099 0.02494 0.01691 -0.0178 0.0488 0.7037
12.500 1.1963 0.02644 0.01913 -0.0240 0.0482 1.0000
12.750 1.2145 0.02721 0.01996 -0.0221 0.0480 1.0000
13.000 1.2322 0.02803 0.02086 -0.0202 0.0477 1.0000
13.250 1.2491 0.02888 0.02179 -0.0182 0.0474 1.0000
13.500 1.2653 0.02977 0.02275 -0.0162 0.0472 1.0000
13.750 1.2804 0.03068 0.02374 -0.0140 0.0469 1.0000
14.000 1.2941 0.03150 0.02463 -0.0117 0.0466 1.0000
14.250 1.3070 0.03234 0.02553 -0.0094 0.0463 1.0000
14.500 1.3193 0.03325 0.02651 -0.0070 0.0461 1.0000
14.750 1.3307 0.03417 0.02749 -0.0046 0.0458 1.0000
15.000 1.3411 0.03507 0.02847 -0.0022 0.0456 1.0000
15.250 1.3510 0.03607 0.02953 0.0002 0.0454 1.0000
15.500 1.3601 0.03709 0.03062 0.0026 0.0453 1.0000
15.750 1.3684 0.03822 0.03182 0.0050 0.0452 1.0000
16.000 1.3763 0.03927 0.03293 0.0073 0.0450 1.0000
16.250 1.3830 0.04051 0.03424 0.0095 0.0449 1.0000
16.500 1.3891 0.04180 0.03560 0.0117 0.0448 1.0000
16.750 1.3946 0.04313 0.03700 0.0139 0.0447 1.0000
17.000 1.3992 0.04455 0.03848 0.0159 0.0446 1.0000
17.250 1.4025 0.04617 0.04019 0.0178 0.0445 1.0000
17.500 1.4042 0.04798 0.04207 0.0196 0.0444 1.0000
17.750 1.4047 0.04993 0.04411 0.0213 0.0443 1.0000
18.000 1.4041 0.05207 0.04633 0.0228 0.0441 1.0000
18.250 1.3907 0.05545 0.04993 0.0242 0.0440 1.0000
18.500 1.3596 0.06090 0.05574 0.0248 0.0437 1.0000
18.750 1.2949 0.07114 0.06653 0.0224 0.0431 1.0000
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Polar data table (+)
Polar graphs
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