N-9 (n9-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: N-9 (n9-il) Reynolds number: 200,000 Max Cl/Cd: 68.1 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n9-il-200000.txt Download as CSV file: xf-n9-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: N-9
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.4812 0.09932 0.09575 -0.0191 1.0000 0.0490
-10.000 -0.4864 0.09515 0.09163 -0.0248 1.0000 0.0494
-9.750 -0.4896 0.09057 0.08711 -0.0305 1.0000 0.0496
-9.500 -0.4904 0.08482 0.08140 -0.0298 1.0000 0.0504
-9.250 -0.4803 0.08269 0.07929 -0.0255 1.0000 0.0516
-9.000 -0.4733 0.07978 0.07639 -0.0257 1.0000 0.0529
-8.750 -0.4682 0.07617 0.07280 -0.0280 1.0000 0.0545
-8.500 -0.4636 0.07189 0.06853 -0.0317 1.0000 0.0565
-8.250 -0.4618 0.06420 0.06045 -0.0446 1.0000 0.0611
-8.000 -0.4603 0.05775 0.05396 -0.0455 1.0000 0.0622
-7.750 -0.4488 0.05590 0.05227 -0.0429 1.0000 0.0638
-7.500 -0.4375 0.05419 0.05059 -0.0413 1.0000 0.0667
-7.250 -0.4255 0.05136 0.04694 -0.0441 1.0000 0.0741
-7.000 -0.4194 0.04483 0.04062 -0.0436 1.0000 0.0760
-6.750 -0.4068 0.03341 0.02823 -0.0432 1.0000 0.0485
-6.500 -0.3915 0.02907 0.02353 -0.0424 1.0000 0.0464
-6.250 -0.3721 0.02590 0.01984 -0.0413 1.0000 0.0473
-6.000 -0.3504 0.02338 0.01683 -0.0402 1.0000 0.0484
-5.750 -0.3273 0.02148 0.01449 -0.0393 1.0000 0.0493
-5.500 -0.3034 0.01921 0.01197 -0.0389 1.0000 0.0520
-5.250 -0.2805 0.01859 0.01130 -0.0382 1.0000 0.0556
-5.000 -0.2559 0.01762 0.01012 -0.0376 1.0000 0.0585
-4.750 -0.2290 0.01680 0.00910 -0.0374 0.9996 0.0622
-4.500 -0.1874 0.01588 0.00821 -0.0404 0.9955 0.0681
-4.250 -0.1454 0.01529 0.00748 -0.0432 0.9903 0.0746
-4.000 -0.1037 0.01445 0.00673 -0.0461 0.9855 0.0838
-3.750 -0.0625 0.01367 0.00600 -0.0489 0.9786 0.0978
-3.500 -0.0181 0.01266 0.00511 -0.0524 0.9727 0.1304
-3.250 0.0231 0.01191 0.00471 -0.0553 0.9633 0.1887
-3.000 0.0656 0.01084 0.00438 -0.0588 0.9567 0.3618
-2.750 0.0966 0.00933 0.00428 -0.0591 0.9480 0.7259
-2.500 0.1566 0.00877 0.00411 -0.0641 0.9446 1.0000
-2.250 0.1973 0.00860 0.00378 -0.0664 0.9320 1.0000
-2.000 0.2339 0.00850 0.00355 -0.0679 0.9186 1.0000
-1.750 0.2667 0.00847 0.00339 -0.0686 0.9042 1.0000
-1.500 0.2958 0.00850 0.00330 -0.0685 0.8887 1.0000
-1.250 0.3222 0.00856 0.00325 -0.0678 0.8723 1.0000
-1.000 0.3473 0.00862 0.00322 -0.0670 0.8550 1.0000
-0.750 0.3727 0.00866 0.00319 -0.0662 0.8379 1.0000
-0.500 0.3983 0.00870 0.00315 -0.0654 0.8211 1.0000
-0.250 0.4239 0.00873 0.00309 -0.0646 0.8033 1.0000
0.000 0.4494 0.00875 0.00301 -0.0638 0.7837 1.0000
0.250 0.4749 0.00882 0.00296 -0.0629 0.7624 1.0000
0.500 0.5006 0.00894 0.00293 -0.0621 0.7423 1.0000
0.750 0.5262 0.00909 0.00300 -0.0614 0.7222 1.0000
1.000 0.5519 0.00928 0.00308 -0.0607 0.7022 1.0000
1.250 0.5775 0.00948 0.00318 -0.0601 0.6821 1.0000
1.500 0.6028 0.00964 0.00330 -0.0594 0.6589 1.0000
1.750 0.6282 0.00982 0.00340 -0.0587 0.6355 1.0000
2.000 0.6533 0.00999 0.00352 -0.0580 0.6087 1.0000
2.250 0.6781 0.01017 0.00362 -0.0573 0.5770 1.0000
2.500 0.7026 0.01038 0.00372 -0.0565 0.5375 1.0000
2.750 0.7266 0.01067 0.00386 -0.0557 0.4952 1.0000
3.000 0.7504 0.01104 0.00407 -0.0549 0.4582 1.0000
3.250 0.7741 0.01146 0.00432 -0.0542 0.4235 1.0000
3.500 0.7980 0.01189 0.00462 -0.0535 0.3940 1.0000
3.750 0.8222 0.01231 0.00494 -0.0530 0.3683 1.0000
4.000 0.8462 0.01273 0.00531 -0.0524 0.3413 1.0000
4.250 0.8698 0.01318 0.00567 -0.0518 0.3043 1.0000
4.500 0.8918 0.01381 0.00606 -0.0509 0.2394 1.0000
4.750 0.9119 0.01477 0.00661 -0.0499 0.1749 1.0000
5.000 0.9332 0.01560 0.00730 -0.0491 0.1547 1.0000
5.250 0.9544 0.01645 0.00805 -0.0482 0.1432 1.0000
5.500 0.9770 0.01713 0.00879 -0.0474 0.1353 1.0000
5.750 0.9976 0.01811 0.00971 -0.0465 0.1281 1.0000
6.000 1.0199 0.01883 0.01050 -0.0458 0.1200 1.0000
6.250 1.0411 0.01974 0.01142 -0.0450 0.1115 1.0000
6.500 1.0620 0.02065 0.01229 -0.0443 0.1029 1.0000
6.750 1.0843 0.02131 0.01311 -0.0436 0.0956 1.0000
7.000 1.1043 0.02240 0.01415 -0.0429 0.0877 1.0000
7.250 1.1268 0.02255 0.01455 -0.0422 0.0788 1.0000
7.500 1.1475 0.02308 0.01519 -0.0414 0.0678 1.0000
7.750 1.1675 0.02366 0.01580 -0.0405 0.0548 1.0000
8.000 1.1854 0.02463 0.01680 -0.0392 0.0454 1.0000
8.250 1.1997 0.02609 0.01829 -0.0376 0.0400 1.0000
8.500 1.2158 0.02737 0.01970 -0.0361 0.0364 1.0000
8.750 1.2295 0.02883 0.02119 -0.0345 0.0338 1.0000
9.000 1.2418 0.03090 0.02339 -0.0326 0.0317 1.0000
9.250 1.2559 0.03231 0.02501 -0.0311 0.0297 1.0000
9.500 1.2683 0.03390 0.02674 -0.0294 0.0281 1.0000
9.750 1.2788 0.03566 0.02860 -0.0276 0.0270 1.0000
10.000 1.2869 0.03789 0.03096 -0.0256 0.0261 1.0000
10.250 1.2914 0.04158 0.03490 -0.0236 0.0253 1.0000
10.500 1.2910 0.04440 0.03802 -0.0210 0.0250 1.0000
10.750 1.2874 0.04705 0.04097 -0.0184 0.0248 1.0000
11.000 1.2803 0.04998 0.04420 -0.0161 0.0245 1.0000
11.250 1.2703 0.05330 0.04782 -0.0145 0.0243 1.0000
11.500 1.2576 0.05710 0.05190 -0.0135 0.0241 1.0000
11.750 1.2423 0.06150 0.05657 -0.0135 0.0239 1.0000
12.000 1.2247 0.06659 0.06192 -0.0144 0.0238 1.0000
12.250 1.2048 0.07246 0.06803 -0.0165 0.0239 1.0000
12.500 1.1827 0.07918 0.07497 -0.0195 0.0240 1.0000
12.750 1.1587 0.08680 0.08280 -0.0236 0.0242 1.0000
13.000 1.1333 0.09537 0.09155 -0.0285 0.0245 1.0000
13.250 1.1065 0.10495 0.10127 -0.0342 0.0249 1.0000
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Polar data table (+)
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