FX 60-157 AIRFOIL (fx60157-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: FX 60-157 AIRFOIL (fx60157-il) Reynolds number: 50,000 Max Cl/Cd: 28.8 at α=9.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx60157-il-50000.txt Download as CSV file: xf-fx60157-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: FX 60-157 AIRFOIL
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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.3897 0.11891 0.11231 -0.0189 1.0000 0.3598
-9.250 -0.5394 0.09705 0.09087 -0.0412 1.0000 0.1712
-9.000 -0.5758 0.08905 0.08291 -0.0437 1.0000 0.1497
-8.750 -0.5713 0.07812 0.07243 -0.0424 1.0000 0.1415
-8.500 -0.5990 0.07417 0.06846 -0.0402 1.0000 0.1367
-8.250 -0.6248 0.07005 0.06432 -0.0381 1.0000 0.1336
-8.000 -0.7368 0.07069 0.06394 -0.0372 1.0000 0.1242
-7.750 -0.7370 0.06636 0.05948 -0.0357 1.0000 0.1212
-7.500 -0.7413 0.06194 0.05472 -0.0344 1.0000 0.1182
-7.250 -0.7415 0.05757 0.04984 -0.0334 1.0000 0.1162
-7.000 -0.7336 0.05396 0.04583 -0.0324 1.0000 0.1172
-6.750 -0.7215 0.05064 0.04209 -0.0316 1.0000 0.1192
-6.500 -0.7061 0.04758 0.03849 -0.0310 1.0000 0.1226
-6.250 -0.6872 0.04475 0.03499 -0.0305 1.0000 0.1259
-6.000 -0.6680 0.04224 0.03256 -0.0295 1.0000 0.1317
-5.750 -0.6462 0.04007 0.03001 -0.0285 1.0000 0.1353
-5.500 -0.6242 0.03824 0.02779 -0.0274 1.0000 0.1409
-5.250 -0.6041 0.03671 0.02633 -0.0259 1.0000 0.1487
-5.000 -0.5830 0.03548 0.02492 -0.0238 1.0000 0.1556
-4.750 -0.5645 0.03449 0.02408 -0.0212 1.0000 0.1659
-4.500 -0.5468 0.03362 0.02329 -0.0188 1.0000 0.1823
-4.250 -0.5310 0.03254 0.02247 -0.0164 1.0000 0.2049
-4.000 -0.5151 0.03072 0.02132 -0.0152 1.0000 0.2611
-3.750 -0.5286 0.03028 0.02384 -0.0034 1.0000 0.5835
-3.500 -0.5338 0.03419 0.02752 0.0092 1.0000 0.6996
-3.250 -0.5335 0.03626 0.02937 0.0182 1.0000 0.7427
-3.000 -0.5304 0.03769 0.03063 0.0263 1.0000 0.7736
-2.750 -0.5229 0.03876 0.03152 0.0333 1.0000 0.8047
-2.500 -0.2890 0.04477 0.03635 0.0115 1.0000 0.9276
-2.250 -0.2373 0.04427 0.03554 0.0053 1.0000 0.9490
-2.000 -0.1897 0.04380 0.03481 -0.0007 1.0000 0.9660
-1.750 -0.1424 0.04353 0.03429 -0.0069 1.0000 0.9819
-1.500 -0.0911 0.04332 0.03387 -0.0141 1.0000 0.9953
-1.250 -0.0700 0.04323 0.03365 -0.0157 1.0000 1.0000
-1.000 -0.0679 0.04308 0.03345 -0.0136 1.0000 1.0000
-0.750 -0.0657 0.04298 0.03330 -0.0116 1.0000 1.0000
-0.500 -0.0637 0.04292 0.03319 -0.0095 1.0000 1.0000
-0.250 -0.0617 0.04289 0.03313 -0.0075 1.0000 1.0000
0.000 -0.0599 0.04291 0.03311 -0.0054 1.0000 1.0000
0.250 -0.0581 0.04296 0.03312 -0.0033 1.0000 1.0000
0.500 -0.0566 0.04303 0.03317 -0.0013 1.0000 1.0000
0.750 -0.0552 0.04313 0.03325 0.0008 1.0000 1.0000
1.000 -0.0540 0.04326 0.03336 0.0028 1.0000 1.0000
1.250 -0.0529 0.04340 0.03348 0.0049 1.0000 1.0000
1.500 -0.0520 0.04356 0.03362 0.0069 1.0000 1.0000
1.750 0.0304 0.04587 0.03586 -0.0061 0.9702 1.0000
2.000 0.0727 0.04669 0.03664 -0.0110 0.9461 1.0000
2.250 0.1148 0.04769 0.03761 -0.0156 0.9259 1.0000
2.500 0.1424 0.04811 0.03802 -0.0175 0.9049 1.0000
2.750 0.1687 0.04855 0.03846 -0.0189 0.8853 1.0000
3.000 0.1965 0.04900 0.03890 -0.0204 0.8665 1.0000
3.250 0.2215 0.04943 0.03935 -0.0213 0.8492 1.0000
3.500 0.2435 0.04984 0.03977 -0.0217 0.8327 1.0000
3.750 0.2663 0.05019 0.04014 -0.0221 0.8161 1.0000
4.000 0.2940 0.05044 0.04042 -0.0229 0.7988 1.0000
4.250 0.3230 0.05057 0.04059 -0.0236 0.7817 1.0000
4.500 0.3378 0.05070 0.04074 -0.0224 0.7641 1.0000
4.750 0.3543 0.05087 0.04096 -0.0214 0.7477 1.0000
5.000 0.3736 0.05103 0.04116 -0.0207 0.7319 1.0000
5.250 0.3930 0.05120 0.04140 -0.0200 0.7155 1.0000
5.500 0.4119 0.05143 0.04168 -0.0193 0.6979 1.0000
5.750 0.4349 0.05167 0.04199 -0.0192 0.6802 1.0000
6.000 0.4628 0.05179 0.04220 -0.0196 0.6619 1.0000
6.250 0.4940 0.05183 0.04234 -0.0202 0.6442 1.0000
6.500 0.5258 0.05194 0.04256 -0.0210 0.6274 1.0000
6.750 0.5586 0.05201 0.04275 -0.0217 0.6107 1.0000
7.000 0.5916 0.05207 0.04295 -0.0225 0.5941 1.0000
7.250 0.6301 0.05153 0.04257 -0.0231 0.5775 1.0000
7.500 0.7464 0.04406 0.03538 -0.0263 0.5696 1.0000
7.750 0.7489 0.04594 0.03734 -0.0248 0.5473 1.0000
8.000 0.8385 0.04138 0.03301 -0.0279 0.5313 1.0000
8.250 0.9714 0.03632 0.02803 -0.0373 0.5026 1.0000
8.500 0.9842 0.03716 0.02897 -0.0354 0.4799 1.0000
8.750 1.0456 0.03669 0.02844 -0.0392 0.4511 1.0000
9.000 1.0613 0.03783 0.02968 -0.0379 0.4307 1.0000
9.250 1.0892 0.03889 0.03082 -0.0383 0.4100 1.0000
9.500 1.1401 0.03959 0.03140 -0.0415 0.3868 1.0000
9.750 1.1526 0.04098 0.03292 -0.0399 0.3688 1.0000
10.000 1.1664 0.04226 0.03431 -0.0385 0.3504 1.0000
10.250 1.1826 0.04364 0.03578 -0.0375 0.3325 1.0000
10.500 1.2002 0.04530 0.03755 -0.0368 0.3164 1.0000
10.750 1.2177 0.04705 0.03941 -0.0361 0.3011 1.0000
11.000 1.2332 0.04868 0.04113 -0.0352 0.2852 1.0000
11.250 1.2449 0.05012 0.04263 -0.0337 0.2687 1.0000
11.500 1.2529 0.05145 0.04401 -0.0318 0.2526 1.0000
11.750 1.2613 0.05293 0.04553 -0.0301 0.2369 1.0000
12.000 1.2714 0.05462 0.04725 -0.0287 0.2216 1.0000
12.250 1.2807 0.05648 0.04915 -0.0274 0.2068 1.0000
12.500 1.2512 0.06021 0.05330 -0.0232 0.2023 1.0000
13.000 1.2610 0.06387 0.05693 -0.0204 0.1750 1.0000
13.250 1.2280 0.06869 0.06213 -0.0178 0.1726 1.0000
13.500 1.1917 0.07458 0.06834 -0.0164 0.1714 1.0000
13.750 1.1454 0.08264 0.07668 -0.0167 0.1736 1.0000
14.000 1.0977 0.09243 0.08664 -0.0191 0.1761 1.0000
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Polar data table (+)
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