NACA 66 (p51htip-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 66 (p51htip-il) Reynolds number: 500,000 Max Cl/Cd: 72.93 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-p51htip-il-500000.txt Download as CSV file: xf-p51htip-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 66
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
------ -------- --------- --------- -------- -------- --------
-10.750 -0.4820 0.08633 0.08415 -0.0541 0.9777 0.0226
-10.500 -0.4817 0.07801 0.07582 -0.0632 0.9643 0.0229
-10.250 -0.5021 0.06870 0.06638 -0.0720 0.9440 0.0226
-10.000 -0.5248 0.06409 0.06164 -0.0729 0.9281 0.0227
-9.750 -0.5488 0.06028 0.05768 -0.0716 0.9153 0.0227
-9.500 -0.5676 0.05740 0.05466 -0.0688 0.9047 0.0227
-9.250 -0.5811 0.05479 0.05190 -0.0658 0.8964 0.0231
-9.000 -0.5894 0.05163 0.04855 -0.0633 0.8900 0.0234
-8.750 -0.5918 0.04846 0.04514 -0.0609 0.8844 0.0249
-8.500 -0.5878 0.04750 0.04372 -0.0574 0.8787 0.0265
-8.250 -0.5861 0.04498 0.04088 -0.0547 0.8744 0.0266
-8.000 -0.5797 0.04222 0.03785 -0.0527 0.8704 0.0266
-7.000 -0.5445 0.02219 0.01609 -0.0437 0.8569 0.0214
-6.750 -0.5209 0.02069 0.01443 -0.0432 0.8539 0.0218
-6.500 -0.4962 0.01953 0.01316 -0.0429 0.8512 0.0224
-6.250 -0.4713 0.01868 0.01221 -0.0427 0.8488 0.0234
-6.000 -0.4459 0.01788 0.01129 -0.0424 0.8466 0.0246
-5.750 -0.4202 0.01706 0.01035 -0.0420 0.8446 0.0255
-5.500 -0.3941 0.01646 0.00965 -0.0418 0.8426 0.0263
-5.250 -0.3691 0.01508 0.00820 -0.0415 0.8404 0.0274
-5.000 -0.3460 0.01413 0.00724 -0.0410 0.8381 0.0287
-4.750 -0.3214 0.01371 0.00683 -0.0407 0.8359 0.0308
-4.500 -0.2968 0.01330 0.00638 -0.0403 0.8339 0.0327
-4.250 -0.2724 0.01290 0.00593 -0.0398 0.8319 0.0342
-4.000 -0.2492 0.01240 0.00536 -0.0391 0.8300 0.0357
-3.750 -0.2266 0.01187 0.00481 -0.0383 0.8283 0.0391
-3.500 -0.2011 0.01159 0.00453 -0.0382 0.8264 0.0425
-3.250 -0.1752 0.01137 0.00428 -0.0380 0.8245 0.0452
-3.000 -0.1504 0.01102 0.00394 -0.0376 0.8226 0.0537
-2.750 -0.1291 0.01032 0.00363 -0.0368 0.8207 0.1436
-2.500 -0.1231 0.00836 0.00314 -0.0339 0.8184 0.5201
-2.250 -0.1065 0.00774 0.00321 -0.0317 0.8166 0.7140
-2.000 -0.0803 0.00778 0.00328 -0.0313 0.8152 0.7493
-1.750 -0.0540 0.00790 0.00343 -0.0308 0.8139 0.7754
-1.500 -0.0282 0.00808 0.00367 -0.0302 0.8124 0.7997
-1.250 -0.0019 0.00825 0.00387 -0.0297 0.8106 0.8151
-1.000 0.0240 0.00843 0.00409 -0.0291 0.8087 0.8292
-0.750 0.0503 0.00861 0.00427 -0.0286 0.8070 0.8406
-0.500 0.0774 0.00870 0.00438 -0.0284 0.8054 0.8462
-0.250 0.1050 0.00875 0.00442 -0.0284 0.8037 0.8509
0.000 0.1330 0.00878 0.00441 -0.0286 0.8021 0.8554
0.250 0.1613 0.00877 0.00440 -0.0288 0.8005 0.8579
0.500 0.1895 0.00882 0.00444 -0.0290 0.7989 0.8602
0.750 0.2171 0.00885 0.00450 -0.0292 0.7963 0.8627
1.000 0.2450 0.00886 0.00454 -0.0295 0.7935 0.8654
1.250 0.2731 0.00885 0.00453 -0.0297 0.7907 0.8685
1.500 0.3016 0.00882 0.00449 -0.0300 0.7881 0.8714
1.750 0.3301 0.00876 0.00445 -0.0302 0.7855 0.8735
2.000 0.3575 0.00875 0.00448 -0.0302 0.7817 0.8759
2.250 0.3849 0.00858 0.00433 -0.0300 0.7747 0.8786
2.500 0.4126 0.00833 0.00407 -0.0297 0.7654 0.8817
2.750 0.4409 0.00809 0.00379 -0.0295 0.7558 0.8851
3.000 0.4673 0.00781 0.00350 -0.0289 0.7405 0.8879
3.250 0.4923 0.00760 0.00332 -0.0281 0.7209 0.8907
3.500 0.5182 0.00751 0.00327 -0.0277 0.7032 0.8940
3.750 0.5433 0.00745 0.00320 -0.0271 0.6742 0.8979
4.000 0.5576 0.00784 0.00311 -0.0244 0.5518 0.9026
4.250 0.5549 0.00930 0.00373 -0.0191 0.3678 0.9092
4.500 0.5515 0.01086 0.00438 -0.0141 0.1735 0.9174
4.750 0.5544 0.01202 0.00498 -0.0099 0.0551 0.9249
5.000 0.5724 0.01241 0.00535 -0.0083 0.0441 0.9314
5.250 0.5874 0.01284 0.00581 -0.0059 0.0386 0.9383
5.500 0.6042 0.01318 0.00621 -0.0040 0.0362 0.9462
5.750 0.6197 0.01350 0.00659 -0.0018 0.0339 0.9552
6.000 0.6390 0.01392 0.00705 -0.0006 0.0317 0.9640
6.250 0.6578 0.01487 0.00804 0.0004 0.0292 0.9742
6.500 0.6864 0.01566 0.00890 -0.0006 0.0279 0.9806
6.750 0.7169 0.01641 0.00970 -0.0020 0.0268 0.9862
7.000 0.7486 0.01720 0.01054 -0.0035 0.0254 0.9908
7.250 0.7802 0.01794 0.01128 -0.0052 0.0238 0.9963
7.500 0.8003 0.01879 0.01213 -0.0045 0.0229 1.0000
8.000 0.8697 0.02230 0.01581 -0.0079 0.0214 1.0000
8.250 0.8858 0.02308 0.01668 -0.0061 0.0210 1.0000
8.500 0.9105 0.02446 0.01821 -0.0058 0.0206 1.0000
8.750 0.9325 0.02571 0.01961 -0.0052 0.0199 1.0000
9.000 0.9536 0.02714 0.02122 -0.0045 0.0192 1.0000
9.250 0.9742 0.02928 0.02358 -0.0037 0.0190 1.0000
9.500 0.9906 0.03205 0.02665 -0.0023 0.0190 1.0000
9.750 0.9984 0.03617 0.03120 0.0003 0.0197 1.0000
10.000 0.9978 0.04080 0.03620 0.0034 0.0207 1.0000
13.500 0.7323 0.12068 0.11840 0.0015 0.0233 1.0000
13.750 0.7273 0.12725 0.12495 -0.0019 0.0232 1.0000
14.000 0.7310 0.13048 0.12819 -0.0034 0.0227 1.0000
14.250 0.7331 0.13422 0.13193 -0.0049 0.0224 1.0000
15.000 0.5769 0.14458 0.14256 -0.0134 0.0228 1.0000
15.250 0.5815 0.14699 0.14497 -0.0146 0.0224 1.0000
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Polar data table (+)
Polar graphs
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