OAF102 AIRFOIL (oaf102-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: OAF102 AIRFOIL (oaf102-il) Reynolds number: 200,000 Max Cl/Cd: 82.22 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-oaf102-il-200000.txt Download as CSV file: xf-oaf102-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: OAF102 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.5526 0.07181 0.06843 -0.0411 1.0000 0.0408
-9.000 -0.5711 0.06190 0.05855 -0.0498 1.0000 0.0396
-8.750 -0.5795 0.04031 0.03611 -0.0838 1.0000 0.0375
-8.500 -0.5568 0.03333 0.02838 -0.0917 1.0000 0.0378
-8.250 -0.5288 0.02923 0.02368 -0.0957 1.0000 0.0390
-8.000 -0.4993 0.02732 0.02125 -0.0979 1.0000 0.0410
-7.750 -0.4731 0.02405 0.01780 -0.0999 1.0000 0.0436
-7.500 -0.4487 0.02268 0.01628 -0.1003 1.0000 0.0460
-7.250 -0.4270 0.02149 0.01491 -0.1000 1.0000 0.0486
-7.000 -0.4076 0.02057 0.01376 -0.0993 1.0000 0.0515
-6.750 -0.3886 0.01936 0.01259 -0.0989 1.0000 0.0552
-6.500 -0.3541 0.01842 0.01150 -0.1009 0.9975 0.0604
-6.250 -0.3126 0.01717 0.01021 -0.1044 0.9945 0.0684
-6.000 -0.2714 0.01611 0.00909 -0.1077 0.9904 0.0784
-5.750 -0.2293 0.01525 0.00824 -0.1111 0.9867 0.0934
-5.500 -0.1858 0.01443 0.00741 -0.1148 0.9840 0.1164
-5.250 -0.1449 0.01376 0.00684 -0.1180 0.9789 0.1472
-5.000 -0.1042 0.01339 0.00650 -0.1210 0.9736 0.1795
-4.750 -0.0651 0.01311 0.00623 -0.1235 0.9678 0.2076
-4.500 -0.0287 0.01294 0.00602 -0.1254 0.9598 0.2326
-4.250 0.0047 0.01282 0.00592 -0.1266 0.9511 0.2540
-4.000 0.0363 0.01279 0.00583 -0.1272 0.9424 0.2745
-3.750 0.0646 0.01276 0.00580 -0.1272 0.9315 0.2926
-3.500 0.0922 0.01275 0.00579 -0.1269 0.9217 0.3094
-3.250 0.1190 0.01275 0.00577 -0.1264 0.9127 0.3256
-3.000 0.1463 0.01274 0.00576 -0.1261 0.9025 0.3415
-2.750 0.1730 0.01273 0.00573 -0.1256 0.8937 0.3572
-2.500 0.2000 0.01271 0.00570 -0.1252 0.8844 0.3730
-2.250 0.2277 0.01270 0.00569 -0.1249 0.8757 0.3889
-2.000 0.2550 0.01269 0.00566 -0.1246 0.8679 0.4048
-1.750 0.2833 0.01269 0.00567 -0.1245 0.8597 0.4210
-1.500 0.3109 0.01268 0.00566 -0.1242 0.8525 0.4370
-1.250 0.3395 0.01269 0.00569 -0.1243 0.8447 0.4533
-1.000 0.3676 0.01268 0.00569 -0.1241 0.8379 0.4696
-0.750 0.3962 0.01269 0.00573 -0.1242 0.8305 0.4858
-0.500 0.4245 0.01268 0.00574 -0.1241 0.8234 0.5015
-0.250 0.4531 0.01268 0.00578 -0.1241 0.8165 0.5169
0.000 0.4818 0.01267 0.00581 -0.1242 0.8093 0.5327
0.250 0.5102 0.01268 0.00585 -0.1241 0.8028 0.5499
0.500 0.5389 0.01266 0.00592 -0.1242 0.7949 0.5680
0.750 0.5668 0.01264 0.00597 -0.1240 0.7886 0.5870
1.000 0.5957 0.01262 0.00607 -0.1242 0.7805 0.6088
1.250 0.6235 0.01258 0.00613 -0.1239 0.7747 0.6346
1.500 0.6522 0.01253 0.00630 -0.1241 0.7664 0.6704
1.750 0.6785 0.01237 0.00638 -0.1234 0.7602 0.7284
2.000 0.6943 0.01179 0.00627 -0.1201 0.7522 1.0000
2.250 0.7242 0.01194 0.00635 -0.1204 0.7456 1.0000
2.500 0.7540 0.01212 0.00654 -0.1208 0.7376 1.0000
2.750 0.7831 0.01227 0.00666 -0.1209 0.7307 1.0000
3.000 0.8123 0.01245 0.00689 -0.1212 0.7222 1.0000
3.250 0.8402 0.01249 0.00690 -0.1207 0.7108 1.0000
3.500 0.8668 0.01232 0.00668 -0.1198 0.6904 1.0000
3.750 0.8931 0.01216 0.00641 -0.1187 0.6684 1.0000
4.000 0.9206 0.01212 0.00640 -0.1182 0.6467 1.0000
4.250 0.9482 0.01216 0.00644 -0.1179 0.6274 1.0000
4.500 0.9756 0.01221 0.00650 -0.1174 0.6039 1.0000
4.750 1.0027 0.01230 0.00657 -0.1170 0.5730 1.0000
5.000 1.0286 0.01251 0.00666 -0.1163 0.5229 1.0000
5.250 1.0495 0.01349 0.00697 -0.1151 0.3914 1.0000
5.500 1.0562 0.01716 0.00886 -0.1133 0.1182 1.0000
5.750 1.0761 0.01864 0.01010 -0.1123 0.0847 1.0000
6.000 1.0974 0.01978 0.01120 -0.1114 0.0717 1.0000
6.250 1.1175 0.02100 0.01243 -0.1103 0.0641 1.0000
6.500 1.1376 0.02216 0.01356 -0.1091 0.0583 1.0000
6.750 1.1564 0.02354 0.01498 -0.1077 0.0538 1.0000
7.000 1.1774 0.02461 0.01610 -0.1066 0.0501 1.0000
7.250 1.1972 0.02596 0.01741 -0.1054 0.0472 1.0000
7.500 1.2174 0.02806 0.01951 -0.1041 0.0446 1.0000
7.750 1.2402 0.02909 0.02068 -0.1032 0.0422 1.0000
8.000 1.2631 0.03050 0.02220 -0.1023 0.0402 1.0000
8.250 1.2860 0.03209 0.02387 -0.1015 0.0386 1.0000
8.500 1.3094 0.03404 0.02586 -0.1009 0.0372 1.0000
8.750 1.3335 0.03781 0.02979 -0.1005 0.0356 1.0000
9.000 1.3517 0.03916 0.03147 -0.0989 0.0344 1.0000
9.250 1.3700 0.04176 0.03442 -0.0974 0.0337 1.0000
9.500 1.3853 0.04494 0.03798 -0.0956 0.0334 1.0000
9.750 1.3965 0.04867 0.04210 -0.0934 0.0333 1.0000
10.000 1.4024 0.05290 0.04674 -0.0908 0.0337 1.0000
10.250 1.4032 0.05723 0.05145 -0.0880 0.0340 1.0000
10.500 1.3992 0.06154 0.05611 -0.0849 0.0342 1.0000
10.750 1.3904 0.06568 0.06057 -0.0815 0.0343 1.0000
11.000 1.3765 0.06983 0.06499 -0.0780 0.0346 1.0000
11.250 1.3584 0.07406 0.06942 -0.0744 0.0349 1.0000
11.500 1.2536 0.07205 0.06754 -0.0588 0.0361 1.0000
11.750 1.2294 0.07485 0.07057 -0.0557 0.0364 1.0000
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Polar data table (+)
Polar graphs
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