N-9 (n9-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: N-9 (n9-il) Reynolds number: 100,000 Max Cl/Cd: 53.28 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n9-il-100000.txt Download as CSV file: xf-n9-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.3633 0.10240 0.09758 -0.0182 1.0000 0.0950
-10.500 -0.3644 0.09899 0.09420 -0.0193 1.0000 0.0983
-10.250 -0.4742 0.10504 0.10001 -0.0131 1.0000 0.0909
-10.000 -0.4657 0.10170 0.09668 -0.0129 1.0000 0.0943
-9.750 -0.4641 0.09838 0.09341 -0.0147 1.0000 0.0977
-9.500 -0.4752 0.09553 0.09066 -0.0198 1.0000 0.1009
-9.250 -0.4925 0.09182 0.08701 -0.0330 1.0000 0.1020
-9.000 -0.4752 0.08698 0.08224 -0.0245 1.0000 0.1041
-8.750 -0.4637 0.08384 0.07912 -0.0222 1.0000 0.1073
-8.500 -0.4597 0.08021 0.07551 -0.0247 1.0000 0.1115
-8.250 -0.4681 0.07455 0.06973 -0.0379 1.0000 0.1171
-8.000 -0.4546 0.07128 0.06659 -0.0325 1.0000 0.1196
-7.750 -0.4451 0.06835 0.06367 -0.0318 1.0000 0.1250
-7.500 -0.4411 0.06365 0.05888 -0.0363 1.0000 0.1334
-7.250 -0.4330 0.06033 0.05538 -0.0390 1.0000 0.1461
-7.000 -0.4208 0.05784 0.05306 -0.0349 1.0000 0.1514
-6.750 -0.4107 0.05445 0.04959 -0.0359 1.0000 0.1631
-6.500 -0.3993 0.05151 0.04657 -0.0360 1.0000 0.1766
-6.250 -0.3657 0.03709 0.03011 -0.0434 1.0000 0.0883
-6.000 -0.3461 0.03403 0.02672 -0.0425 1.0000 0.0875
-5.750 -0.3258 0.03077 0.02312 -0.0417 1.0000 0.0870
-5.500 -0.3037 0.02791 0.01986 -0.0410 1.0000 0.0865
-5.250 -0.2807 0.02560 0.01709 -0.0403 1.0000 0.0893
-5.000 -0.2581 0.02397 0.01543 -0.0397 1.0000 0.0939
-4.750 -0.2334 0.02246 0.01359 -0.0390 1.0000 0.0969
-4.500 -0.2086 0.02135 0.01213 -0.0383 1.0000 0.1032
-4.250 -0.1839 0.02006 0.01085 -0.0379 1.0000 0.1095
-4.000 -0.1587 0.01922 0.00980 -0.0374 1.0000 0.1170
-3.750 -0.1333 0.01828 0.00894 -0.0373 1.0000 0.1301
-3.500 -0.1077 0.01736 0.00821 -0.0372 1.0000 0.1466
-3.250 -0.0834 0.01684 0.00794 -0.0370 1.0000 0.1783
-3.000 -0.0585 0.01632 0.00769 -0.0371 1.0000 0.2349
-2.750 -0.0308 0.01400 0.00761 -0.0371 1.0000 0.6948
-2.500 -0.0160 0.01380 0.00773 -0.0345 1.0000 1.0000
-2.250 0.0339 0.01412 0.00768 -0.0394 0.9904 1.0000
-2.000 0.0909 0.01430 0.00758 -0.0454 0.9769 1.0000
-1.750 0.1443 0.01435 0.00742 -0.0506 0.9633 1.0000
-1.500 0.1916 0.01441 0.00733 -0.0546 0.9515 1.0000
-1.250 0.2385 0.01443 0.00724 -0.0585 0.9413 1.0000
-1.000 0.2879 0.01436 0.00709 -0.0627 0.9322 1.0000
-0.750 0.3298 0.01432 0.00699 -0.0653 0.9203 1.0000
-0.500 0.3705 0.01428 0.00690 -0.0676 0.9084 1.0000
-0.250 0.4091 0.01423 0.00682 -0.0693 0.8959 1.0000
0.000 0.4439 0.01421 0.00678 -0.0702 0.8820 1.0000
0.250 0.4751 0.01422 0.00679 -0.0703 0.8666 1.0000
0.500 0.5040 0.01422 0.00679 -0.0699 0.8501 1.0000
0.750 0.5318 0.01419 0.00675 -0.0691 0.8329 1.0000
1.000 0.5586 0.01407 0.00662 -0.0680 0.8136 1.0000
1.250 0.5832 0.01395 0.00649 -0.0663 0.7902 1.0000
1.500 0.6086 0.01383 0.00632 -0.0648 0.7661 1.0000
1.750 0.6345 0.01378 0.00618 -0.0634 0.7414 1.0000
2.000 0.6594 0.01388 0.00621 -0.0620 0.7138 1.0000
2.250 0.6837 0.01408 0.00630 -0.0606 0.6837 1.0000
2.500 0.7079 0.01431 0.00647 -0.0593 0.6534 1.0000
2.750 0.7318 0.01456 0.00664 -0.0581 0.6227 1.0000
3.000 0.7553 0.01478 0.00685 -0.0569 0.5896 1.0000
3.250 0.7788 0.01501 0.00703 -0.0557 0.5565 1.0000
3.500 0.8021 0.01524 0.00718 -0.0545 0.5230 1.0000
3.750 0.8256 0.01555 0.00741 -0.0535 0.4922 1.0000
4.000 0.8487 0.01593 0.00772 -0.0525 0.4616 1.0000
4.250 0.8710 0.01638 0.00809 -0.0513 0.4270 1.0000
4.500 0.8918 0.01697 0.00852 -0.0500 0.3870 1.0000
4.750 0.9109 0.01772 0.00911 -0.0485 0.3338 1.0000
5.000 0.9278 0.01879 0.00986 -0.0467 0.2681 1.0000
5.250 0.9469 0.01982 0.01060 -0.0454 0.2256 1.0000
5.500 0.9681 0.02079 0.01146 -0.0443 0.2023 1.0000
5.750 0.9896 0.02181 0.01238 -0.0434 0.1866 1.0000
6.000 1.0115 0.02288 0.01342 -0.0425 0.1734 1.0000
6.250 1.0342 0.02407 0.01462 -0.0418 0.1625 1.0000
6.500 1.0566 0.02540 0.01593 -0.0411 0.1513 1.0000
6.750 1.0784 0.02685 0.01737 -0.0404 0.1391 1.0000
7.000 1.0986 0.02838 0.01890 -0.0397 0.1252 1.0000
7.250 1.1186 0.03020 0.02079 -0.0389 0.1117 1.0000
7.500 1.1372 0.03225 0.02294 -0.0379 0.0979 1.0000
7.750 1.1550 0.03457 0.02536 -0.0368 0.0853 1.0000
8.000 1.1724 0.03645 0.02727 -0.0357 0.0750 1.0000
8.250 1.1859 0.03854 0.02986 -0.0336 0.0676 1.0000
8.500 1.2038 0.04099 0.03230 -0.0329 0.0627 1.0000
8.750 1.2116 0.04399 0.03592 -0.0304 0.0594 1.0000
9.000 1.2206 0.04637 0.03866 -0.0284 0.0557 1.0000
9.250 1.2326 0.04841 0.04079 -0.0271 0.0528 1.0000
9.500 1.2386 0.05277 0.04534 -0.0257 0.0513 1.0000
9.750 1.2325 0.05664 0.04969 -0.0228 0.0508 1.0000
10.000 1.2216 0.06063 0.05410 -0.0199 0.0506 1.0000
10.250 1.2057 0.06450 0.05831 -0.0169 0.0505 1.0000
10.500 1.1862 0.06843 0.06250 -0.0143 0.0506 1.0000
10.750 1.1654 0.07280 0.06710 -0.0130 0.0507 1.0000
11.000 1.1438 0.07774 0.07224 -0.0130 0.0509 1.0000
11.250 1.1224 0.08336 0.07803 -0.0142 0.0511 1.0000
11.500 0.7718 0.12047 0.11580 -0.0298 0.0814 1.0000
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Polar data table (+)
Polar graphs
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