GIII BL0 AIRFOIL (giiia-il) Xfoil prediction polar at RE=200,000 Ncrit=9
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Airfoil: GIII BL0 AIRFOIL (giiia-il) Reynolds number: 200,000 Max Cl/Cd: 49.74 at α=7.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-giiia-il-200000.txt Download as CSV file: xf-giiia-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GIII BL0 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.4840 0.09949 0.09618 -0.0153 1.0000 0.0652
-10.500 -0.4952 0.09379 0.09052 -0.0180 1.0000 0.0673
-8.750 -0.7787 0.04810 0.04325 -0.0188 1.0000 0.0346
-8.500 -0.7814 0.04391 0.03875 -0.0156 1.0000 0.0340
-8.250 -0.7806 0.04048 0.03495 -0.0121 1.0000 0.0342
-8.000 -0.7774 0.03734 0.03143 -0.0085 1.0000 0.0343
-7.750 -0.7724 0.03411 0.02782 -0.0050 1.0000 0.0341
-7.500 -0.7643 0.03127 0.02460 -0.0016 1.0000 0.0341
-7.250 -0.7529 0.02898 0.02194 0.0015 1.0000 0.0345
-7.000 -0.7392 0.02727 0.01990 0.0043 1.0000 0.0351
-6.750 -0.7241 0.02458 0.01694 0.0064 1.0000 0.0369
-6.500 -0.7073 0.02331 0.01565 0.0082 1.0000 0.0393
-6.250 -0.6890 0.02212 0.01430 0.0101 1.0000 0.0413
-6.000 -0.6700 0.02097 0.01302 0.0120 1.0000 0.0435
-5.750 -0.6516 0.02032 0.01220 0.0139 1.0000 0.0459
-5.500 -0.6340 0.01872 0.01066 0.0156 1.0000 0.0504
-5.250 -0.6165 0.01797 0.00987 0.0176 1.0000 0.0547
-5.000 -0.6000 0.01705 0.00892 0.0197 1.0000 0.0598
-4.750 -0.5832 0.01642 0.00830 0.0217 1.0000 0.0680
-4.500 -0.5668 0.01565 0.00757 0.0237 1.0000 0.0782
-4.250 -0.5492 0.01502 0.00699 0.0255 1.0000 0.0967
-4.000 -0.5326 0.01422 0.00647 0.0273 1.0000 0.1358
-3.750 -0.5150 0.01355 0.00612 0.0289 1.0000 0.1980
-3.500 -0.4973 0.01291 0.00578 0.0303 1.0000 0.2621
-3.250 -0.4835 0.01178 0.00538 0.0322 1.0000 0.4026
-3.000 -0.4660 0.01065 0.00541 0.0342 0.9978 0.6605
-2.750 -0.4304 0.01048 0.00564 0.0332 0.9935 0.7641
-2.500 -0.3937 0.01067 0.00597 0.0323 0.9879 0.8299
-2.250 -0.3524 0.01105 0.00634 0.0304 0.9834 0.8706
-2.000 -0.3133 0.01144 0.00667 0.0288 0.9773 0.8964
-1.750 -0.2679 0.01184 0.00698 0.0259 0.9724 0.9137
-1.500 -0.2110 0.01226 0.00730 0.0205 0.9695 0.9230
-1.250 -0.1562 0.01248 0.00742 0.0156 0.9618 0.9305
-1.000 -0.0836 0.01261 0.00746 0.0070 0.9573 0.9332
-0.750 -0.0262 0.01263 0.00742 0.0015 0.9490 0.9397
-0.500 0.0337 0.01257 0.00730 -0.0044 0.9407 0.9445
-0.250 0.0807 0.01244 0.00714 -0.0078 0.9273 0.9498
0.000 0.1140 0.01227 0.00693 -0.0086 0.9113 0.9571
0.250 0.1527 0.01206 0.00668 -0.0105 0.8923 0.9614
0.500 0.1815 0.01192 0.00650 -0.0103 0.8700 0.9696
0.750 0.2193 0.01169 0.00621 -0.0121 0.8447 0.9739
1.000 0.2537 0.01153 0.00599 -0.0133 0.8178 0.9804
1.250 0.2900 0.01135 0.00574 -0.0150 0.7863 0.9858
1.500 0.3268 0.01119 0.00547 -0.0168 0.7470 0.9911
1.750 0.3633 0.01109 0.00519 -0.0186 0.7040 0.9959
2.000 0.4004 0.01108 0.00497 -0.0206 0.6612 1.0000
2.250 0.4231 0.01117 0.00488 -0.0198 0.6233 1.0000
2.500 0.4459 0.01129 0.00484 -0.0191 0.5845 1.0000
2.750 0.4685 0.01146 0.00484 -0.0183 0.5406 1.0000
3.000 0.4906 0.01170 0.00485 -0.0175 0.4844 1.0000
3.250 0.5118 0.01213 0.00489 -0.0166 0.3927 1.0000
3.500 0.5328 0.01274 0.00509 -0.0158 0.3198 1.0000
3.750 0.5548 0.01317 0.00531 -0.0151 0.2899 1.0000
4.000 0.5772 0.01353 0.00557 -0.0144 0.2745 1.0000
4.250 0.5994 0.01392 0.00586 -0.0136 0.2644 1.0000
4.500 0.6221 0.01423 0.00617 -0.0129 0.2559 1.0000
4.750 0.6447 0.01463 0.00653 -0.0121 0.2496 1.0000
5.000 0.6676 0.01494 0.00688 -0.0114 0.2442 1.0000
5.250 0.6905 0.01534 0.00727 -0.0106 0.2397 1.0000
5.500 0.7136 0.01582 0.00776 -0.0099 0.2359 1.0000
5.750 0.7360 0.01611 0.00817 -0.0090 0.2304 1.0000
6.000 0.7572 0.01651 0.00853 -0.0081 0.2211 1.0000
6.250 0.7772 0.01669 0.00875 -0.0069 0.2095 1.0000
6.500 0.7973 0.01685 0.00902 -0.0057 0.1990 1.0000
6.750 0.8171 0.01706 0.00931 -0.0045 0.1886 1.0000
7.000 0.8363 0.01722 0.00953 -0.0031 0.1770 1.0000
7.250 0.8557 0.01721 0.00966 -0.0017 0.1620 1.0000
7.500 0.8724 0.01754 0.00958 0.0000 0.0673 1.0000
7.750 0.8825 0.01908 0.01099 0.0029 0.0456 1.0000
8.000 0.8926 0.02035 0.01235 0.0058 0.0396 1.0000
8.250 0.9055 0.02125 0.01338 0.0084 0.0362 1.0000
8.500 0.9154 0.02231 0.01449 0.0112 0.0330 1.0000
8.750 0.9186 0.02405 0.01622 0.0149 0.0308 1.0000
9.000 0.9306 0.02507 0.01736 0.0175 0.0297 1.0000
9.250 0.9420 0.02629 0.01868 0.0202 0.0286 1.0000
9.500 0.9538 0.02766 0.02014 0.0226 0.0276 1.0000
9.750 0.9664 0.02917 0.02174 0.0249 0.0268 1.0000
10.000 0.9783 0.03067 0.02335 0.0271 0.0258 1.0000
10.250 0.9890 0.03226 0.02501 0.0293 0.0247 1.0000
10.500 1.0005 0.03509 0.02791 0.0309 0.0235 1.0000
11.000 1.0121 0.04083 0.03416 0.0356 0.0230 1.0000
11.250 1.0132 0.04318 0.03677 0.0384 0.0229 1.0000
11.500 1.0101 0.04583 0.03966 0.0414 0.0229 1.0000
11.750 1.0032 0.04892 0.04297 0.0444 0.0230 1.0000
12.000 0.9964 0.05196 0.04624 0.0469 0.0231 1.0000
12.250 0.9880 0.05433 0.04881 0.0489 0.0232 1.0000
12.500 0.9783 0.05690 0.05157 0.0502 0.0234 1.0000
12.750 0.9668 0.05980 0.05468 0.0507 0.0235 1.0000
13.000 0.9526 0.06331 0.05840 0.0502 0.0237 1.0000
13.250 0.9347 0.06781 0.06315 0.0481 0.0241 1.0000
13.500 0.9105 0.07410 0.06971 0.0442 0.0245 1.0000
13.750 0.8767 0.08308 0.07897 0.0373 0.0250 1.0000
14.000 0.8353 0.09537 0.09152 0.0278 0.0256 1.0000
14.250 0.7058 0.09185 0.08812 0.0371 0.0256 1.0000
14.500 0.6402 0.10809 0.10455 0.0275 0.0271 1.0000
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Polar data table (+)
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