NACA 8-H-12 AIRFOIL (n8h12-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 8-H-12 AIRFOIL (n8h12-il) Reynolds number: 100,000 Max Cl/Cd: 45.61 at α=9.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n8h12-il-100000.txt Download as CSV file: xf-n8h12-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 8-H-12 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4099 0.09687 0.09291 -0.0190 0.9130 0.1038
-8.750 -0.4375 0.09421 0.09025 -0.0269 0.8937 0.1048
-8.500 -0.4627 0.09220 0.08811 -0.0313 0.8793 0.1052
-8.250 -0.4006 0.08674 0.08275 -0.0203 0.8712 0.1107
-8.000 -0.4017 0.08371 0.07974 -0.0219 0.8586 0.1144
-7.750 -0.4199 0.08064 0.07663 -0.0265 0.8467 0.1185
-7.500 -0.4398 0.07691 0.07269 -0.0313 0.8370 0.1213
-7.250 -0.4173 0.07353 0.06943 -0.0274 0.8278 0.1246
-7.000 -0.4102 0.07080 0.06664 -0.0280 0.8179 0.1319
-6.750 -0.4165 0.06713 0.06276 -0.0298 0.8105 0.1383
-6.500 -0.3992 0.06463 0.06029 -0.0286 0.8006 0.1463
-6.250 -0.3960 0.06140 0.05689 -0.0289 0.7935 0.1559
-6.000 -0.3887 0.05877 0.05403 -0.0297 0.7846 0.1697
-5.750 -0.3729 0.05609 0.05140 -0.0276 0.7778 0.1764
-5.500 -0.3600 0.05338 0.04858 -0.0275 0.7690 0.1908
-5.250 -0.3495 0.05105 0.04610 -0.0258 0.7632 0.2066
-5.000 -0.3358 0.04913 0.04402 -0.0257 0.7541 0.2346
-4.750 -0.3028 0.03878 0.03169 -0.0266 0.7500 0.1032
-4.500 -0.2800 0.03541 0.02753 -0.0247 0.7436 0.0893
-4.250 -0.2571 0.03267 0.02453 -0.0239 0.7362 0.0869
-4.000 -0.2353 0.03069 0.02213 -0.0219 0.7312 0.0859
-3.750 -0.2078 0.02935 0.02044 -0.0218 0.7223 0.0880
-3.500 -0.1829 0.02804 0.01875 -0.0203 0.7169 0.0892
-3.250 -0.1544 0.02681 0.01727 -0.0202 0.7097 0.0907
-3.000 -0.1277 0.02538 0.01586 -0.0199 0.7034 0.0954
-2.750 -0.1026 0.02456 0.01489 -0.0186 0.6990 0.1021
-2.500 -0.0733 0.02366 0.01406 -0.0191 0.6904 0.1088
-2.250 -0.0489 0.02291 0.01329 -0.0179 0.6855 0.1213
-2.000 -0.0252 0.02218 0.01276 -0.0172 0.6796 0.1437
-1.750 0.1465 0.01863 0.01192 -0.0408 0.6728 1.0000
-1.500 0.1675 0.01866 0.01169 -0.0395 0.6683 1.0000
-1.250 0.1933 0.01899 0.01188 -0.0400 0.6614 1.0000
-1.000 0.2166 0.01917 0.01188 -0.0395 0.6558 1.0000
-0.750 0.2382 0.01928 0.01180 -0.0383 0.6521 1.0000
-0.500 0.2647 0.01982 0.01227 -0.0391 0.6458 1.0000
-0.250 0.2889 0.02016 0.01250 -0.0390 0.6406 1.0000
0.000 0.3109 0.02031 0.01249 -0.0378 0.6369 1.0000
0.250 0.3359 0.02086 0.01298 -0.0381 0.6315 1.0000
0.500 0.3609 0.02145 0.01352 -0.0385 0.6259 1.0000
0.750 0.3839 0.02177 0.01375 -0.0378 0.6223 1.0000
1.000 0.4061 0.02200 0.01386 -0.0366 0.6196 1.0000
1.250 0.4322 0.02318 0.01510 -0.0383 0.6129 1.0000
1.500 0.4554 0.02372 0.01561 -0.0380 0.6083 1.0000
1.750 0.4779 0.02403 0.01585 -0.0370 0.6053 1.0000
2.000 0.5003 0.02437 0.01611 -0.0359 0.6030 1.0000
2.250 0.5234 0.02632 0.01820 -0.0384 0.5955 1.0000
2.500 0.5453 0.02692 0.01878 -0.0379 0.5916 1.0000
2.750 0.5676 0.02724 0.01907 -0.0367 0.5890 1.0000
3.000 0.5865 0.02912 0.02103 -0.0379 0.5837 1.0000
3.250 0.6041 0.03072 0.02268 -0.0384 0.5779 1.0000
3.500 0.6254 0.03129 0.02325 -0.0375 0.5748 1.0000
3.750 0.6483 0.03151 0.02346 -0.0361 0.5726 1.0000
4.000 0.6486 0.03566 0.02776 -0.0383 0.5633 1.0000
4.250 0.6678 0.03644 0.02856 -0.0373 0.5597 1.0000
4.500 0.6940 0.03627 0.02839 -0.0357 0.5574 1.0000
4.750 0.6745 0.04164 0.03385 -0.0369 0.5470 1.0000
5.000 0.6998 0.04173 0.03398 -0.0356 0.5434 1.0000
5.250 0.7331 0.04092 0.03320 -0.0338 0.5413 1.0000
5.500 0.7002 0.04658 0.03888 -0.0338 0.5294 1.0000
5.750 0.7386 0.04568 0.03805 -0.0325 0.5264 1.0000
6.000 0.7005 0.05058 0.04291 -0.0303 0.5159 1.0000
6.250 0.7436 0.04981 0.04226 -0.0301 0.5116 1.0000
6.500 0.7895 0.04829 0.04085 -0.0288 0.5089 1.0000
6.750 0.7527 0.05291 0.04542 -0.0265 0.4965 1.0000
7.000 0.7373 0.05586 0.04836 -0.0244 0.4866 1.0000
7.250 0.7776 0.05494 0.04757 -0.0236 0.4810 1.0000
7.500 0.8441 0.05196 0.04481 -0.0234 0.4779 1.0000
8.500 1.0739 0.03247 0.02585 -0.0129 0.4495 1.0000
8.750 1.1131 0.02914 0.02256 -0.0107 0.4344 1.0000
9.000 1.1458 0.02673 0.02015 -0.0088 0.4167 1.0000
9.250 1.1612 0.02632 0.01997 -0.0070 0.3947 1.0000
9.500 1.1757 0.02578 0.01958 -0.0050 0.3636 1.0000
9.750 1.1756 0.02641 0.02022 -0.0025 0.3071 1.0000
10.000 1.1535 0.02868 0.02190 0.0011 0.2409 1.0000
10.250 1.1264 0.03214 0.02501 0.0038 0.2055 1.0000
10.500 1.1007 0.03627 0.02883 0.0053 0.1761 1.0000
10.750 1.0777 0.04055 0.03283 0.0062 0.1513 1.0000
11.000 1.0607 0.04441 0.03642 0.0071 0.1321 1.0000
11.250 1.0507 0.04754 0.03927 0.0082 0.1157 1.0000
11.500 1.0513 0.04955 0.04110 0.0098 0.1020 1.0000
11.750 1.0598 0.05083 0.04223 0.0116 0.0914 1.0000
12.000 1.0728 0.05192 0.04330 0.0133 0.0831 1.0000
12.250 1.0956 0.05227 0.04349 0.0157 0.0764 1.0000
12.500 1.1167 0.05320 0.04452 0.0174 0.0721 1.0000
12.750 1.1486 0.05372 0.04492 0.0194 0.0681 1.0000
13.000 1.1747 0.05536 0.04669 0.0209 0.0652 1.0000
13.250 1.1836 0.05765 0.04922 0.0216 0.0631 1.0000
13.500 1.1956 0.06004 0.05184 0.0223 0.0614 1.0000
13.750 1.2052 0.06277 0.05478 0.0230 0.0604 1.0000
14.000 1.2101 0.06585 0.05808 0.0234 0.0597 1.0000
14.250 1.2095 0.06936 0.06182 0.0236 0.0593 1.0000
14.500 1.2027 0.07329 0.06601 0.0234 0.0591 1.0000
14.750 1.1889 0.07780 0.07079 0.0228 0.0591 1.0000
15.000 1.1704 0.08288 0.07614 0.0215 0.0592 1.0000
15.250 1.1435 0.08900 0.08254 0.0193 0.0596 1.0000
15.500 1.1086 0.09664 0.09049 0.0159 0.0606 1.0000
15.750 1.0717 0.10556 0.09967 0.0113 0.0617 1.0000
16.000 1.0309 0.11616 0.11048 0.0053 0.0629 1.0000
16.250 0.9892 0.12839 0.12285 -0.0018 0.0643 1.0000
16.500 0.9560 0.14048 0.13500 -0.0085 0.0658 1.0000
16.750 0.9459 0.14828 0.14282 -0.0116 0.0667 1.0000
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