ULTIMATE (ultimate-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: ULTIMATE (ultimate-il) Reynolds number: 50,000 Max Cl/Cd: 28.36 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ultimate-il-50000.txt Download as CSV file: xf-ultimate-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: ULTIMATE
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.750 -0.6324 0.14259 0.13546 0.0225 1.0000 0.2869
-11.500 -0.6215 0.13710 0.12996 0.0216 1.0000 0.2884
-11.250 -0.7809 0.09491 0.08810 -0.0193 1.0000 0.1267
-11.000 -0.7707 0.08994 0.08312 -0.0192 1.0000 0.1247
-10.750 -0.7925 0.08296 0.07611 -0.0222 1.0000 0.1221
-10.500 -0.8296 0.07636 0.06942 -0.0241 1.0000 0.1194
-10.250 -0.8733 0.07124 0.06407 -0.0222 1.0000 0.1168
-10.000 -0.9292 0.06686 0.05899 -0.0175 1.0000 0.1126
-9.750 -0.9326 0.06249 0.05434 -0.0155 1.0000 0.1122
-9.500 -0.9329 0.05829 0.04981 -0.0133 1.0000 0.1117
-9.250 -0.9311 0.05433 0.04544 -0.0110 1.0000 0.1115
-9.000 -0.9273 0.05070 0.04132 -0.0084 1.0000 0.1120
-8.750 -0.9167 0.04715 0.03745 -0.0065 1.0000 0.1145
-8.500 -0.8976 0.04434 0.03459 -0.0056 1.0000 0.1197
-8.250 -0.8822 0.04145 0.03121 -0.0037 1.0000 0.1236
-8.000 -0.8619 0.03851 0.02788 -0.0024 1.0000 0.1282
-7.750 -0.8388 0.03625 0.02554 -0.0016 1.0000 0.1384
-7.500 -0.8123 0.03385 0.02314 -0.0012 1.0000 0.1504
-7.250 -0.7871 0.03167 0.02091 -0.0003 1.0000 0.1687
-7.000 -0.7653 0.02946 0.01899 0.0009 1.0000 0.1994
-6.750 -0.7548 0.02720 0.01740 0.0040 1.0000 0.2578
-6.500 -0.7605 0.02559 0.01672 0.0105 1.0000 0.3602
-6.250 -0.7602 0.02514 0.01667 0.0167 1.0000 0.4423
-6.000 -0.7519 0.02507 0.01674 0.0216 1.0000 0.5002
-5.750 -0.7360 0.02547 0.01732 0.0258 1.0000 0.5495
-5.500 -0.7158 0.02648 0.01846 0.0303 1.0000 0.5984
-5.250 -0.6954 0.02712 0.01908 0.0342 1.0000 0.6387
-5.000 -0.6706 0.02739 0.01925 0.0368 1.0000 0.6695
-4.750 -0.6449 0.02750 0.01923 0.0388 1.0000 0.6970
-4.500 -0.6123 0.02775 0.01932 0.0399 1.0000 0.7224
-4.250 -0.5872 0.02764 0.01907 0.0415 1.0000 0.7479
-4.000 -0.5522 0.02780 0.01904 0.0416 1.0000 0.7727
-3.750 -0.5097 0.02803 0.01906 0.0403 1.0000 0.7967
-3.500 -0.4790 0.02788 0.01875 0.0403 1.0000 0.8224
-3.250 -0.4059 0.02841 0.01897 0.0333 1.0000 0.8439
-3.000 -0.3601 0.02824 0.01862 0.0300 1.0000 0.8686
-2.750 -0.2834 0.02829 0.01840 0.0211 1.0000 0.8885
-2.500 -0.2291 0.02794 0.01789 0.0155 1.0000 0.9113
-2.250 -0.1667 0.02747 0.01727 0.0081 1.0000 0.9323
-2.000 -0.1078 0.02692 0.01661 0.0009 1.0000 0.9542
-1.750 -0.0446 0.02622 0.01583 -0.0075 1.0000 0.9747
-1.500 0.0208 0.02535 0.01490 -0.0168 1.0000 0.9951
-1.250 0.0434 0.02468 0.01423 -0.0186 1.0000 1.0000
-1.000 0.0422 0.02436 0.01394 -0.0161 1.0000 1.0000
-0.750 0.0361 0.02419 0.01379 -0.0126 1.0000 1.0000
-0.500 0.0259 0.02413 0.01375 -0.0085 1.0000 1.0000
-0.250 0.0119 0.02415 0.01377 -0.0039 1.0000 1.0000
0.000 -0.0047 0.02423 0.01385 0.0011 1.0000 1.0000
0.250 -0.0222 0.02433 0.01394 0.0062 1.0000 1.0000
0.500 -0.0371 0.02449 0.01406 0.0110 1.0000 1.0000
0.750 -0.0460 0.02476 0.01428 0.0148 1.0000 1.0000
1.000 -0.0483 0.02516 0.01461 0.0177 1.0000 1.0000
1.250 0.0051 0.02599 0.01547 0.0104 0.9858 1.0000
1.500 0.0623 0.02678 0.01633 0.0028 0.9677 1.0000
1.750 0.1166 0.02749 0.01714 -0.0040 0.9496 1.0000
2.000 0.1720 0.02812 0.01789 -0.0107 0.9323 1.0000
2.250 0.2198 0.02868 0.01856 -0.0157 0.9142 1.0000
2.500 0.2641 0.02919 0.01923 -0.0198 0.8956 1.0000
2.750 0.3146 0.02963 0.01985 -0.0248 0.8774 1.0000
3.000 0.3677 0.02988 0.02034 -0.0296 0.8591 1.0000
3.250 0.4111 0.03005 0.02072 -0.0323 0.8397 1.0000
3.500 0.4388 0.03039 0.02120 -0.0322 0.8182 1.0000
3.750 0.4784 0.03027 0.02132 -0.0331 0.7978 1.0000
4.000 0.5057 0.03037 0.02160 -0.0322 0.7767 1.0000
4.250 0.5320 0.03039 0.02181 -0.0308 0.7550 1.0000
4.500 0.5659 0.02987 0.02152 -0.0295 0.7349 1.0000
4.750 0.5839 0.03009 0.02192 -0.0268 0.7104 1.0000
5.000 0.6101 0.02924 0.02125 -0.0233 0.6824 1.0000
5.250 0.6336 0.02668 0.01867 -0.0159 0.6354 1.0000
5.500 0.6510 0.02511 0.01704 -0.0097 0.5857 1.0000
5.750 0.6670 0.02430 0.01618 -0.0048 0.5338 1.0000
6.000 0.6746 0.02379 0.01545 0.0012 0.4372 1.0000
6.250 0.6668 0.02610 0.01601 0.0080 0.2426 1.0000
6.500 0.6790 0.02858 0.01780 0.0106 0.1867 1.0000
6.750 0.6998 0.03051 0.01948 0.0122 0.1612 1.0000
7.000 0.7250 0.03247 0.02128 0.0132 0.1452 1.0000
7.250 0.7492 0.03433 0.02321 0.0143 0.1331 1.0000
7.500 0.7748 0.03672 0.02586 0.0153 0.1270 1.0000
7.750 0.7974 0.03904 0.02836 0.0164 0.1209 1.0000
8.000 0.8174 0.04184 0.03133 0.0175 0.1160 1.0000
8.250 0.8327 0.04480 0.03481 0.0193 0.1139 1.0000
8.500 0.8452 0.04821 0.03868 0.0212 0.1134 1.0000
8.750 0.8534 0.05187 0.04279 0.0232 0.1134 1.0000
9.000 0.8569 0.05568 0.04700 0.0254 0.1134 1.0000
9.250 0.8547 0.05969 0.05140 0.0276 0.1135 1.0000
9.500 0.8481 0.06388 0.05591 0.0297 0.1141 1.0000
9.750 0.8387 0.06825 0.06053 0.0315 0.1150 1.0000
10.000 0.8296 0.07279 0.06522 0.0330 0.1161 1.0000
10.250 0.8334 0.07777 0.07028 0.0336 0.1176 1.0000
10.500 0.7254 0.08632 0.07910 0.0320 0.1281 1.0000
10.750 0.7077 0.09360 0.08639 0.0288 0.1329 1.0000
11.000 0.5076 0.10384 0.09677 0.0228 0.1551 1.0000
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Polar data table (+)
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