S6063 7.05% (s6063-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: S6063 7.05% (s6063-il) Reynolds number: 1,000,000 Max Cl/Cd: 81.25 at α=2.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-s6063-il-1000000.txt Download as CSV file: xf-s6063-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: S6063 7.05%
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
------ -------- --------- --------- -------- -------- --------
-5.250 -0.3965 0.00655 0.00299 -0.0380 0.9919 0.0129
-5.000 -0.3661 0.00541 0.00167 -0.0392 0.9867 0.0140
-4.750 -0.3364 0.00448 0.00054 -0.0399 0.9804 0.0153
-4.500 -0.3048 0.00495 0.00096 -0.0401 0.9713 0.0172
-4.000 -0.2637 0.01004 0.00505 -0.0393 0.9732 0.0088
-3.750 -0.2358 0.00910 0.00396 -0.0391 0.9642 0.0084
-3.500 -0.2093 0.00850 0.00327 -0.0386 0.9548 0.0089
-3.250 -0.1840 0.00809 0.00278 -0.0379 0.9442 0.0098
-3.000 -0.1584 0.00776 0.00238 -0.0372 0.9333 0.0105
-2.750 -0.1325 0.00751 0.00205 -0.0366 0.9226 0.0110
-2.500 -0.1064 0.00727 0.00172 -0.0361 0.9119 0.0119
-2.250 -0.0800 0.00708 0.00144 -0.0355 0.9011 0.0137
-2.000 -0.0534 0.00674 0.00120 -0.0351 0.8896 0.0474
-1.750 -0.0277 0.00615 0.00102 -0.0349 0.8780 0.1720
-1.500 -0.0017 0.00569 0.00091 -0.0347 0.8660 0.2869
-1.250 0.0240 0.00520 0.00082 -0.0345 0.8534 0.4244
-1.000 0.0493 0.00467 0.00078 -0.0342 0.8403 0.5858
-0.750 0.0744 0.00430 0.00076 -0.0336 0.8267 0.7090
-0.500 0.0995 0.00408 0.00075 -0.0328 0.8123 0.7943
-0.250 0.1245 0.00395 0.00075 -0.0320 0.7970 0.8507
0.000 0.1487 0.00388 0.00075 -0.0308 0.7804 0.9006
0.250 0.1732 0.00386 0.00075 -0.0297 0.7625 0.9431
0.500 0.2056 0.00389 0.00073 -0.0305 0.7423 0.9728
1.000 0.2788 0.00404 0.00069 -0.0343 0.6944 1.0000
1.250 0.3046 0.00415 0.00071 -0.0339 0.6690 1.0000
1.500 0.3306 0.00429 0.00074 -0.0335 0.6401 1.0000
1.750 0.3567 0.00445 0.00077 -0.0331 0.6073 1.0000
2.000 0.3824 0.00471 0.00083 -0.0327 0.5537 1.0000
2.250 0.4079 0.00504 0.00091 -0.0323 0.4897 1.0000
2.500 0.4339 0.00534 0.00101 -0.0321 0.4373 1.0000
2.750 0.4597 0.00569 0.00113 -0.0318 0.3755 1.0000
3.000 0.4844 0.00629 0.00131 -0.0315 0.2778 1.0000
3.250 0.5095 0.00683 0.00152 -0.0313 0.1989 1.0000
3.500 0.5336 0.00762 0.00185 -0.0310 0.0947 1.0000
3.750 0.5579 0.00843 0.00227 -0.0306 0.0160 1.0000
4.000 0.5848 0.00871 0.00257 -0.0303 0.0126 1.0000
4.250 0.6115 0.00903 0.00298 -0.0300 0.0116 1.0000
4.500 0.6378 0.00944 0.00346 -0.0297 0.0112 1.0000
4.750 0.6642 0.00979 0.00387 -0.0294 0.0109 1.0000
5.000 0.6902 0.01023 0.00436 -0.0291 0.0107 1.0000
5.250 0.7160 0.01073 0.00493 -0.0287 0.0105 1.0000
5.500 0.7413 0.01131 0.00558 -0.0283 0.0103 1.0000
5.750 0.7662 0.01195 0.00631 -0.0278 0.0100 1.0000
6.000 0.7906 0.01273 0.00716 -0.0272 0.0096 1.0000
6.250 0.8144 0.01367 0.00819 -0.0265 0.0092 1.0000
6.500 0.8373 0.01495 0.00958 -0.0256 0.0093 1.0000
6.750 0.8593 0.01684 0.01164 -0.0245 0.0098 1.0000
7.250 0.9055 0.02195 0.01730 -0.0219 0.0138 1.0000
7.500 0.9277 0.02337 0.01884 -0.0213 0.0131 1.0000
7.750 0.9488 0.02498 0.02057 -0.0207 0.0125 1.0000
8.000 0.9651 0.02788 0.02367 -0.0198 0.0117 1.0000
8.250 0.9598 0.03645 0.03295 -0.0172 0.0111 1.0000
8.500 0.9701 0.03929 0.03609 -0.0158 0.0111 1.0000
8.750 0.9976 0.03784 0.03475 -0.0151 0.0102 1.0000
9.000 1.0123 0.03961 0.03671 -0.0139 0.0090 1.0000
9.250 1.0171 0.04280 0.04014 -0.0125 0.0087 1.0000
9.500 1.0204 0.04546 0.04298 -0.0112 0.0083 1.0000
9.750 1.0167 0.04870 0.04641 -0.0097 0.0081 1.0000
10.000 1.0051 0.05180 0.04967 -0.0075 0.0080 1.0000
10.250 0.9851 0.05522 0.05324 -0.0057 0.0080 1.0000
10.500 0.9564 0.06110 0.05931 -0.0071 0.0082 1.0000
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Polar data table (+)
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