N-9 (n9-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: N-9 (n9-il) Reynolds number: 500,000 Max Cl/Cd: 84.04 at α=1.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n9-il-500000.txt Download as CSV file: xf-n9-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5413 0.04603 0.04351 -0.0490 1.0000 0.0184
-8.750 -0.5500 0.03531 0.03223 -0.0501 1.0000 0.0176
-8.500 -0.5501 0.02803 0.02421 -0.0492 1.0000 0.0177
-8.250 -0.5401 0.02408 0.01967 -0.0476 1.0000 0.0180
-8.000 -0.5284 0.02050 0.01567 -0.0459 1.0000 0.0187
-7.750 -0.5103 0.01957 0.01465 -0.0445 1.0000 0.0193
-7.500 -0.4829 0.01865 0.01359 -0.0448 0.9990 0.0202
-7.250 -0.4493 0.01754 0.01227 -0.0462 0.9971 0.0216
-7.000 -0.4141 0.01697 0.01151 -0.0476 0.9952 0.0231
-6.750 -0.3822 0.01486 0.00916 -0.0489 0.9935 0.0248
-6.500 -0.3487 0.01419 0.00844 -0.0501 0.9906 0.0264
-6.250 -0.3140 0.01356 0.00771 -0.0515 0.9878 0.0285
-6.000 -0.2779 0.01325 0.00731 -0.0531 0.9854 0.0304
-5.750 -0.2447 0.01188 0.00588 -0.0545 0.9830 0.0337
-5.500 -0.2121 0.01140 0.00536 -0.0554 0.9781 0.0364
-5.250 -0.1755 0.01107 0.00498 -0.0571 0.9743 0.0389
-5.000 -0.1401 0.01019 0.00407 -0.0587 0.9698 0.0431
-4.750 -0.1060 0.00981 0.00368 -0.0599 0.9622 0.0469
-4.500 -0.0713 0.00942 0.00324 -0.0611 0.9543 0.0507
-4.250 -0.0356 0.00900 0.00285 -0.0626 0.9457 0.0582
-4.000 -0.0021 0.00869 0.00256 -0.0636 0.9347 0.0716
-3.750 0.0293 0.00844 0.00235 -0.0641 0.9219 0.0943
-3.500 0.0581 0.00829 0.00216 -0.0641 0.9079 0.1093
-3.250 0.0853 0.00815 0.00201 -0.0637 0.8924 0.1237
-3.000 0.1116 0.00794 0.00189 -0.0632 0.8762 0.1570
-2.750 0.1379 0.00770 0.00182 -0.0628 0.8603 0.2288
-2.500 0.1642 0.00740 0.00172 -0.0625 0.8433 0.2981
-2.250 0.1888 0.00672 0.00166 -0.0620 0.8215 0.4980
-2.000 0.2121 0.00614 0.00161 -0.0610 0.7967 0.6796
-1.750 0.2290 0.00557 0.00164 -0.0578 0.7730 0.8937
-1.500 0.2753 0.00559 0.00159 -0.0613 0.7509 0.9906
-1.250 0.3093 0.00572 0.00154 -0.0626 0.7294 1.0000
-1.000 0.3349 0.00586 0.00153 -0.0620 0.7126 1.0000
-0.750 0.3608 0.00598 0.00153 -0.0615 0.6971 1.0000
-0.500 0.3869 0.00609 0.00155 -0.0610 0.6814 1.0000
-0.250 0.4132 0.00620 0.00157 -0.0606 0.6675 1.0000
0.000 0.4398 0.00629 0.00160 -0.0603 0.6539 1.0000
0.250 0.4664 0.00639 0.00163 -0.0599 0.6395 1.0000
0.500 0.4930 0.00649 0.00167 -0.0596 0.6239 1.0000
0.750 0.5196 0.00659 0.00171 -0.0593 0.6050 1.0000
1.000 0.5458 0.00673 0.00175 -0.0589 0.5788 1.0000
1.250 0.5719 0.00689 0.00181 -0.0585 0.5489 1.0000
1.500 0.5975 0.00711 0.00188 -0.0580 0.5068 1.0000
1.750 0.6220 0.00747 0.00199 -0.0575 0.4530 1.0000
2.000 0.6469 0.00784 0.00216 -0.0570 0.4104 1.0000
2.250 0.6723 0.00817 0.00234 -0.0566 0.3794 1.0000
2.500 0.6980 0.00848 0.00253 -0.0563 0.3545 1.0000
2.750 0.7242 0.00873 0.00271 -0.0561 0.3355 1.0000
3.000 0.7503 0.00899 0.00291 -0.0559 0.3170 1.0000
3.250 0.7763 0.00926 0.00311 -0.0557 0.2953 1.0000
3.500 0.8020 0.00957 0.00331 -0.0554 0.2638 1.0000
3.750 0.8252 0.01021 0.00360 -0.0548 0.1892 1.0000
4.000 0.8469 0.01107 0.00411 -0.0541 0.1314 1.0000
4.250 0.8717 0.01151 0.00451 -0.0537 0.1208 1.0000
4.500 0.8966 0.01193 0.00493 -0.0533 0.1138 1.0000
4.750 0.9219 0.01229 0.00532 -0.0530 0.1090 1.0000
5.000 0.9454 0.01287 0.00587 -0.0525 0.1018 1.0000
5.250 0.9716 0.01308 0.00614 -0.0523 0.0974 1.0000
5.500 0.9967 0.01342 0.00649 -0.0521 0.0921 1.0000
5.750 1.0205 0.01391 0.00700 -0.0516 0.0867 1.0000
6.000 1.0467 0.01410 0.00725 -0.0515 0.0826 1.0000
6.250 1.0717 0.01441 0.00757 -0.0512 0.0776 1.0000
6.500 1.0964 0.01476 0.00796 -0.0509 0.0726 1.0000
6.750 1.1223 0.01495 0.00817 -0.0508 0.0663 1.0000
7.000 1.1472 0.01525 0.00849 -0.0506 0.0588 1.0000
7.250 1.1712 0.01566 0.00882 -0.0503 0.0450 1.0000
7.500 1.1912 0.01655 0.00959 -0.0493 0.0290 1.0000
7.750 1.2112 0.01742 0.01047 -0.0483 0.0234 1.0000
8.000 1.2309 0.01829 0.01137 -0.0473 0.0204 1.0000
8.250 1.2491 0.01928 0.01246 -0.0460 0.0187 1.0000
8.500 1.2684 0.02009 0.01337 -0.0449 0.0175 1.0000
8.750 1.2864 0.02100 0.01436 -0.0437 0.0165 1.0000
9.000 1.3016 0.02213 0.01556 -0.0422 0.0155 1.0000
9.250 1.3093 0.02394 0.01749 -0.0397 0.0145 1.0000
9.500 1.3260 0.02478 0.01846 -0.0384 0.0140 1.0000
9.750 1.3397 0.02585 0.01964 -0.0367 0.0134 1.0000
10.000 1.3499 0.02704 0.02095 -0.0346 0.0130 1.0000
10.250 1.3572 0.02829 0.02232 -0.0321 0.0125 1.0000
10.500 1.3632 0.02967 0.02379 -0.0297 0.0122 1.0000
10.750 1.3677 0.03123 0.02546 -0.0274 0.0119 1.0000
11.000 1.3700 0.03306 0.02740 -0.0253 0.0116 1.0000
11.250 1.3691 0.03535 0.02981 -0.0233 0.0113 1.0000
11.500 1.3645 0.03822 0.03283 -0.0213 0.0111 1.0000
11.750 1.3559 0.04179 0.03661 -0.0196 0.0109 1.0000
12.000 1.3523 0.04482 0.03983 -0.0187 0.0108 1.0000
12.250 1.3504 0.04774 0.04293 -0.0184 0.0107 1.0000
12.500 1.3467 0.05107 0.04645 -0.0186 0.0106 1.0000
12.750 1.3413 0.05480 0.05035 -0.0193 0.0104 1.0000
13.000 1.3339 0.05900 0.05473 -0.0203 0.0103 1.0000
13.250 1.3244 0.06364 0.05955 -0.0218 0.0102 1.0000
13.500 1.3133 0.06874 0.06484 -0.0237 0.0102 1.0000
13.750 1.3000 0.07445 0.07071 -0.0261 0.0101 1.0000
14.000 1.2851 0.08070 0.07714 -0.0290 0.0100 1.0000
14.250 1.2688 0.08760 0.08421 -0.0325 0.0100 1.0000
14.500 1.2506 0.09520 0.09198 -0.0365 0.0100 1.0000
14.750 1.2316 0.10328 0.10022 -0.0408 0.0101 1.0000
15.000 1.2109 0.11200 0.10910 -0.0456 0.0101 1.0000
15.250 1.1886 0.12132 0.11857 -0.0507 0.0102 1.0000
15.500 1.1652 0.13132 0.12872 -0.0563 0.0103 1.0000
15.750 1.1381 0.14277 0.14031 -0.0627 0.0105 1.0000
16.000 1.1043 0.15665 0.15430 -0.0703 0.0108 1.0000
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Polar data table (+)
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