USA 35 A AIRFOIL (usa35a-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: USA 35 A AIRFOIL (usa35a-il) Reynolds number: 100,000 Max Cl/Cd: 46.39 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa35a-il-100000-n5.txt Download as CSV file: xf-usa35a-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA 35 A 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 0.0702 0.11069 0.10537 -0.0934 0.8765 0.0816
-10.500 0.0919 0.10773 0.10241 -0.0923 0.8650 0.0827
-10.000 0.0157 0.10834 0.10284 -0.0912 0.8871 0.0820
-9.500 0.0429 0.10297 0.09742 -0.0916 0.8658 0.0840
-9.000 0.0309 0.09162 0.08593 -0.0980 0.8452 0.0622
-8.750 0.0422 0.08913 0.08338 -0.0981 0.8355 0.0616
-8.500 0.0481 0.08625 0.08050 -0.0989 0.8236 0.0608
-8.250 0.0521 0.08289 0.07708 -0.1003 0.8140 0.0601
-8.000 0.0534 0.07945 0.07364 -0.1019 0.8023 0.0599
-7.750 0.0543 0.07577 0.06990 -0.1039 0.7932 0.0601
-7.500 0.0520 0.07200 0.06614 -0.1061 0.7813 0.0603
-7.250 0.0462 0.06774 0.06182 -0.1090 0.7723 0.0602
-7.000 0.0402 0.06206 0.05609 -0.1147 0.7609 0.0599
-6.750 0.0282 0.05289 0.04662 -0.1228 0.7525 0.0591
-6.500 0.0201 0.04528 0.03855 -0.1269 0.7430 0.0591
-6.250 0.0261 0.04064 0.03342 -0.1283 0.7338 0.0596
-6.000 0.0387 0.03697 0.02921 -0.1289 0.7254 0.0609
-5.750 0.0528 0.03382 0.02543 -0.1288 0.7154 0.0626
-5.500 0.0783 0.03296 0.02447 -0.1288 0.7077 0.0636
-5.250 0.0995 0.03178 0.02312 -0.1283 0.6978 0.0648
-5.000 0.1231 0.03024 0.02121 -0.1281 0.6900 0.0661
-4.750 0.1472 0.02868 0.01920 -0.1279 0.6828 0.0677
-4.500 0.1715 0.02786 0.01823 -0.1275 0.6742 0.0692
-4.250 0.1984 0.02722 0.01747 -0.1274 0.6674 0.0713
-4.000 0.2237 0.02641 0.01639 -0.1270 0.6596 0.0738
-3.750 0.2494 0.02573 0.01557 -0.1267 0.6523 0.0757
-3.500 0.2772 0.02515 0.01487 -0.1266 0.6466 0.0778
-3.250 0.3033 0.02465 0.01418 -0.1262 0.6399 0.0809
-3.000 0.3292 0.02422 0.01374 -0.1259 0.6330 0.0836
-2.750 0.3569 0.02378 0.01318 -0.1258 0.6273 0.0869
-2.500 0.3838 0.02339 0.01269 -0.1255 0.6216 0.0904
-2.250 0.4090 0.02311 0.01243 -0.1250 0.6144 0.0941
-2.000 0.4362 0.02276 0.01194 -0.1247 0.6079 0.0983
-1.750 0.4628 0.02247 0.01165 -0.1244 0.6016 0.1028
-1.500 0.4877 0.02226 0.01141 -0.1238 0.5942 0.1077
-1.250 0.5142 0.02201 0.01117 -0.1235 0.5883 0.1133
-1.000 0.5426 0.02178 0.01086 -0.1235 0.5836 0.1199
-0.750 0.5668 0.02171 0.01084 -0.1228 0.5776 0.1276
-0.500 0.5924 0.02159 0.01078 -0.1224 0.5719 0.1370
-0.250 0.6196 0.02146 0.01068 -0.1223 0.5670 0.1505
0.000 0.6482 0.02134 0.01057 -0.1223 0.5627 0.1732
0.250 0.6719 0.02131 0.01072 -0.1217 0.5569 0.2104
0.750 0.7231 0.02093 0.01089 -0.1211 0.5468 0.3736
1.000 0.7501 0.02062 0.01092 -0.1208 0.5427 0.5218
1.250 0.7651 0.02021 0.01129 -0.1177 0.5376 0.7488
1.500 0.8034 0.02011 0.01141 -0.1192 0.5314 1.0000
1.750 0.8292 0.02034 0.01147 -0.1188 0.5263 1.0000
2.000 0.8574 0.02056 0.01147 -0.1188 0.5220 1.0000
2.250 0.8803 0.02091 0.01175 -0.1180 0.5170 1.0000
2.500 0.9025 0.02125 0.01204 -0.1171 0.5113 1.0000
2.750 0.9273 0.02153 0.01220 -0.1166 0.5061 1.0000
3.000 0.9547 0.02176 0.01227 -0.1165 0.5015 1.0000
3.250 0.9771 0.02213 0.01258 -0.1156 0.4961 1.0000
3.500 0.9974 0.02252 0.01296 -0.1145 0.4901 1.0000
3.750 1.0213 0.02283 0.01319 -0.1139 0.4848 1.0000
4.000 1.0483 0.02308 0.01331 -0.1137 0.4805 1.0000
4.250 1.0685 0.02352 0.01374 -0.1126 0.4752 1.0000
4.500 1.0870 0.02397 0.01421 -0.1112 0.4691 1.0000
4.750 1.1099 0.02428 0.01446 -0.1105 0.4639 1.0000
5.000 1.1369 0.02451 0.01455 -0.1103 0.4596 1.0000
5.250 1.1504 0.02511 0.01523 -0.1082 0.4532 1.0000
5.500 1.1685 0.02557 0.01569 -0.1068 0.4475 1.0000
5.750 1.1914 0.02589 0.01594 -0.1062 0.4430 1.0000
6.000 1.2134 0.02628 0.01628 -0.1053 0.4387 1.0000
6.250 1.2235 0.02698 0.01706 -0.1029 0.4326 1.0000
6.500 1.2393 0.02744 0.01752 -0.1011 0.4274 1.0000
6.750 1.2620 0.02771 0.01770 -0.1004 0.4231 1.0000
7.000 1.2700 0.02849 0.01855 -0.0977 0.4171 1.0000
7.250 1.2810 0.02920 0.01928 -0.0955 0.4113 1.0000
7.500 1.2998 0.02957 0.01959 -0.0944 0.4064 1.0000
7.750 1.3099 0.03037 0.02042 -0.0923 0.4006 1.0000
8.000 1.3170 0.03130 0.02140 -0.0899 0.3941 1.0000
8.250 1.3346 0.03172 0.02174 -0.0887 0.3888 1.0000
8.500 1.3414 0.03281 0.02287 -0.0866 0.3828 1.0000
8.750 1.3484 0.03394 0.02406 -0.0846 0.3769 1.0000
9.000 1.3648 0.03456 0.02463 -0.0835 0.3724 1.0000
9.250 1.3776 0.03549 0.02556 -0.0822 0.3679 1.0000
9.500 1.3795 0.03716 0.02734 -0.0802 0.3626 1.0000
9.750 1.3899 0.03828 0.02848 -0.0788 0.3579 1.0000
10.000 1.4103 0.03874 0.02887 -0.0782 0.3541 1.0000
10.250 1.4125 0.04055 0.03077 -0.0764 0.3494 1.0000
10.500 1.4125 0.04259 0.03291 -0.0747 0.3442 1.0000
10.750 1.4235 0.04377 0.03410 -0.0737 0.3398 1.0000
11.000 1.4481 0.04388 0.03411 -0.0733 0.3361 1.0000
11.250 1.4339 0.04724 0.03767 -0.0712 0.3306 1.0000
11.500 1.4326 0.04960 0.04011 -0.0699 0.3254 1.0000
11.750 1.4475 0.05050 0.04099 -0.0692 0.3213 1.0000
12.000 1.4631 0.05138 0.04186 -0.0685 0.3175 1.0000
12.250 1.4375 0.05629 0.04700 -0.0669 0.3113 1.0000
12.500 1.4402 0.05851 0.04928 -0.0661 0.3065 1.0000
12.750 1.4626 0.05865 0.04935 -0.0656 0.3029 1.0000
13.000 1.4381 0.06394 0.05483 -0.0647 0.2967 1.0000
13.250 1.4231 0.06839 0.05941 -0.0641 0.2904 1.0000
13.500 1.4440 0.06859 0.05958 -0.0636 0.2867 1.0000
13.750 1.4053 0.07622 0.06743 -0.0636 0.2790 1.0000
14.000 1.3956 0.08039 0.07168 -0.0636 0.2728 1.0000
14.250 1.4221 0.07978 0.07101 -0.0630 0.2699 1.0000
14.500 1.3465 0.09320 0.08474 -0.0646 0.2582 1.0000
14.750 1.3690 0.09304 0.08456 -0.0641 0.2554 1.0000
15.000 1.3996 0.09175 0.08321 -0.0634 0.2535 1.0000
15.500 1.3127 0.11095 0.10273 -0.0670 0.2358 1.0000
|
Polar data table (+)
Polar graphs
<< Back to USA 35 A AIRFOIL (usa35a-il)