FX 63-120 AIRFOIL (fx63120-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: FX 63-120 AIRFOIL (fx63120-il) Reynolds number: 1,000,000 Max Cl/Cd: 147.56 at α=3° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx63120-il-1000000.txt Download as CSV file: xf-fx63120-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: FX 63-120 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.250 -0.3304 0.13247 0.13075 -0.0351 1.0000 0.0147
-12.000 -0.3280 0.12907 0.12736 -0.0358 1.0000 0.0156
-11.750 -0.6774 0.03636 0.03394 -0.1047 0.9883 0.0135
-11.500 -0.6439 0.03211 0.02957 -0.1159 0.9853 0.0138
-11.250 -0.6071 0.03236 0.02989 -0.1178 0.9846 0.0141
-11.000 -0.5777 0.02810 0.02540 -0.1285 0.9776 0.0143
-10.750 -0.5430 0.02601 0.02314 -0.1344 0.9751 0.0147
-10.500 -0.5074 0.02390 0.02083 -0.1397 0.9734 0.0151
-10.250 -0.4827 0.02197 0.01869 -0.1421 0.9657 0.0156
-10.000 -0.4486 0.02050 0.01700 -0.1453 0.9625 0.0161
-9.750 -0.4211 0.01827 0.01450 -0.1479 0.9547 0.0166
-9.500 -0.3883 0.01744 0.01363 -0.1499 0.9494 0.0171
-9.250 -0.3534 0.01703 0.01318 -0.1518 0.9440 0.0175
-9.000 -0.3151 0.01637 0.01244 -0.1547 0.9384 0.0180
-8.750 -0.2717 0.01568 0.01163 -0.1586 0.9347 0.0188
-8.500 -0.2291 0.01514 0.01096 -0.1621 0.9283 0.0196
-8.250 -0.1870 0.01380 0.00942 -0.1662 0.9202 0.0204
-8.000 -0.1500 0.01328 0.00883 -0.1685 0.9088 0.0210
-7.750 -0.1171 0.01292 0.00838 -0.1699 0.8971 0.0217
-7.500 -0.0864 0.01257 0.00791 -0.1707 0.8863 0.0224
-7.250 -0.0576 0.01225 0.00748 -0.1711 0.8760 0.0232
-7.000 -0.0287 0.01208 0.00720 -0.1713 0.8668 0.0238
-6.750 -0.0003 0.01119 0.00621 -0.1719 0.8577 0.0247
-6.500 0.0284 0.01076 0.00575 -0.1723 0.8496 0.0255
-6.250 0.0571 0.01046 0.00539 -0.1726 0.8413 0.0263
-6.000 0.0859 0.01019 0.00506 -0.1729 0.8337 0.0271
-5.750 0.1147 0.00996 0.00477 -0.1731 0.8256 0.0281
-5.500 0.1435 0.00977 0.00451 -0.1733 0.8176 0.0287
-5.250 0.1731 0.00929 0.00394 -0.1739 0.8083 0.0295
-5.000 0.2031 0.00887 0.00348 -0.1745 0.7990 0.0306
-4.750 0.2323 0.00868 0.00322 -0.1748 0.7902 0.0318
-4.500 0.2620 0.00848 0.00299 -0.1752 0.7829 0.0329
-4.250 0.2915 0.00831 0.00276 -0.1755 0.7755 0.0339
-4.000 0.3211 0.00817 0.00258 -0.1758 0.7687 0.0346
-3.750 0.3515 0.00792 0.00226 -0.1764 0.7614 0.0364
-3.500 0.3811 0.00779 0.00209 -0.1768 0.7549 0.0386
-3.250 0.4107 0.00767 0.00195 -0.1770 0.7478 0.0406
-2.750 0.4703 0.00743 0.00171 -0.1778 0.7345 0.0567
-2.500 0.5015 0.00713 0.00158 -0.1786 0.7277 0.1184
-2.250 0.5345 0.00663 0.00146 -0.1802 0.7213 0.2559
-2.000 0.5653 0.00640 0.00147 -0.1810 0.7144 0.3544
-1.750 0.5941 0.00642 0.00148 -0.1811 0.7080 0.3787
-1.500 0.6232 0.00642 0.00149 -0.1812 0.7013 0.3934
-1.250 0.6518 0.00646 0.00150 -0.1812 0.6943 0.4070
-1.000 0.6807 0.00647 0.00151 -0.1813 0.6878 0.4204
-0.750 0.7096 0.00648 0.00153 -0.1814 0.6807 0.4300
-0.500 0.7381 0.00652 0.00155 -0.1815 0.6739 0.4400
-0.250 0.7670 0.00653 0.00158 -0.1816 0.6665 0.4559
0.000 0.7955 0.00656 0.00161 -0.1816 0.6592 0.4716
0.250 0.8243 0.00657 0.00165 -0.1818 0.6503 0.4870
0.500 0.8526 0.00661 0.00170 -0.1818 0.6398 0.5027
1.000 0.9092 0.00671 0.00182 -0.1818 0.6185 0.5422
1.250 0.9372 0.00678 0.00191 -0.1818 0.6090 0.5639
1.500 0.9653 0.00686 0.00199 -0.1817 0.5984 0.5826
1.750 0.9933 0.00694 0.00207 -0.1817 0.5882 0.5983
2.000 1.0208 0.00705 0.00217 -0.1815 0.5769 0.6117
2.250 1.0481 0.00718 0.00226 -0.1813 0.5641 0.6240
2.500 1.0754 0.00730 0.00237 -0.1812 0.5496 0.6355
2.750 1.1021 0.00747 0.00250 -0.1809 0.5320 0.6457
3.000 1.1288 0.00765 0.00263 -0.1806 0.5143 0.6557
3.250 1.1551 0.00785 0.00278 -0.1803 0.4944 0.6658
3.500 1.1811 0.00808 0.00295 -0.1799 0.4737 0.6754
3.750 1.2071 0.00831 0.00313 -0.1795 0.4565 0.6852
4.000 1.2331 0.00855 0.00332 -0.1792 0.4396 0.6941
4.250 1.2588 0.00881 0.00351 -0.1788 0.4223 0.7036
4.500 1.2844 0.00906 0.00373 -0.1784 0.4055 0.7128
4.750 1.3099 0.00933 0.00395 -0.1779 0.3879 0.7223
5.000 1.3353 0.00959 0.00417 -0.1775 0.3692 0.7324
5.250 1.3603 0.00989 0.00442 -0.1770 0.3487 0.7424
5.500 1.3846 0.01024 0.00469 -0.1764 0.3267 0.7533
5.750 1.4085 0.01062 0.00499 -0.1757 0.3018 0.7649
6.000 1.4313 0.01109 0.00535 -0.1749 0.2744 0.7774
6.250 1.4534 0.01160 0.00575 -0.1740 0.2474 0.7922
6.500 1.4758 0.01205 0.00614 -0.1731 0.2243 0.8111
6.750 1.4979 0.01249 0.00655 -0.1721 0.2050 0.8362
7.000 1.5175 0.01285 0.00695 -0.1706 0.1880 0.8906
7.250 1.5338 0.01316 0.00727 -0.1684 0.1720 1.0000
7.500 1.5553 0.01370 0.00771 -0.1674 0.1527 1.0000
7.750 1.5746 0.01439 0.00825 -0.1661 0.1298 1.0000
8.000 1.5936 0.01506 0.00880 -0.1647 0.1095 1.0000
8.250 1.6117 0.01577 0.00940 -0.1632 0.0934 1.0000
8.500 1.6303 0.01639 0.00996 -0.1617 0.0825 1.0000
8.750 1.6493 0.01695 0.01050 -0.1603 0.0747 1.0000
9.000 1.6656 0.01758 0.01109 -0.1585 0.0672 1.0000
9.250 1.6818 0.01811 0.01162 -0.1565 0.0618 1.0000
9.500 1.6957 0.01875 0.01224 -0.1542 0.0554 1.0000
9.750 1.7086 0.01946 0.01293 -0.1519 0.0496 1.0000
10.000 1.7225 0.02014 0.01361 -0.1497 0.0454 1.0000
10.250 1.7345 0.02094 0.01441 -0.1474 0.0417 1.0000
10.500 1.7474 0.02171 0.01520 -0.1452 0.0390 1.0000
10.750 1.7602 0.02251 0.01603 -0.1431 0.0371 1.0000
11.000 1.7708 0.02348 0.01702 -0.1408 0.0350 1.0000
11.250 1.7804 0.02455 0.01813 -0.1385 0.0330 1.0000
11.500 1.7932 0.02542 0.01907 -0.1367 0.0321 1.0000
11.750 1.8040 0.02647 0.02016 -0.1348 0.0307 1.0000
12.000 1.8128 0.02771 0.02144 -0.1328 0.0293 1.0000
12.250 1.8183 0.02929 0.02305 -0.1306 0.0275 1.0000
12.500 1.8300 0.03040 0.02424 -0.1291 0.0266 1.0000
12.750 1.8400 0.03169 0.02559 -0.1276 0.0255 1.0000
13.000 1.8476 0.03323 0.02718 -0.1259 0.0242 1.0000
13.250 1.8514 0.03519 0.02918 -0.1241 0.0227 1.0000
13.500 1.8597 0.03679 0.03085 -0.1228 0.0218 1.0000
13.750 1.8678 0.03845 0.03257 -0.1216 0.0206 1.0000
14.000 1.8726 0.04049 0.03466 -0.1203 0.0193 1.0000
14.250 1.8741 0.04295 0.03718 -0.1190 0.0181 1.0000
14.500 1.8805 0.04496 0.03927 -0.1180 0.0173 1.0000
14.750 1.8845 0.04730 0.04167 -0.1171 0.0163 1.0000
15.000 1.8854 0.05008 0.04450 -0.1162 0.0154 1.0000
15.250 1.8843 0.05320 0.04770 -0.1155 0.0146 1.0000
15.500 1.8869 0.05596 0.05055 -0.1149 0.0140 1.0000
15.750 1.8876 0.05904 0.05372 -0.1146 0.0134 1.0000
16.000 1.8863 0.06248 0.05724 -0.1143 0.0129 1.0000
16.250 1.8828 0.06634 0.06118 -0.1143 0.0124 1.0000
16.500 1.8747 0.07101 0.06596 -0.1146 0.0119 1.0000
16.750 1.8724 0.07494 0.07000 -0.1149 0.0116 1.0000
17.000 1.8694 0.07907 0.07424 -0.1154 0.0113 1.0000
17.250 1.8649 0.08354 0.07882 -0.1162 0.0111 1.0000
17.500 1.8597 0.08821 0.08360 -0.1172 0.0108 1.0000
17.750 1.8533 0.09320 0.08870 -0.1184 0.0105 1.0000
18.000 1.8456 0.09851 0.09412 -0.1199 0.0103 1.0000
18.250 1.8348 0.10442 0.10014 -0.1218 0.0100 1.0000
18.500 1.8236 0.11056 0.10640 -0.1240 0.0098 1.0000
18.750 1.8095 0.11734 0.11331 -0.1267 0.0096 1.0000
|
Polar data table (+)
Polar graphs
<< Back to FX 63-120 AIRFOIL (fx63120-il)