NACA 66(1)-212 (naca661212-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA 66(1)-212 (naca661212-il) Reynolds number: 100,000 Max Cl/Cd: 32.69 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-naca661212-il-100000-n5.txt Download as CSV file: xf-naca661212-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 66(1)-212
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.5682 0.07481 0.06948 -0.0649 1.0000 0.0301
-10.250 -0.5833 0.07121 0.06588 -0.0644 1.0000 0.0299
-10.000 -0.6028 0.06849 0.06314 -0.0622 1.0000 0.0297
-9.750 -0.6277 0.06689 0.06156 -0.0577 1.0000 0.0296
-9.500 -0.6581 0.06619 0.06087 -0.0511 1.0000 0.0296
-9.250 -0.6667 0.06261 0.05712 -0.0502 0.9937 0.0294
-9.000 -0.6659 0.05811 0.05229 -0.0511 0.9859 0.0294
-8.750 -0.6636 0.05422 0.04793 -0.0510 0.9781 0.0298
-8.500 -0.6590 0.05008 0.04339 -0.0509 0.9715 0.0306
-8.250 -0.6462 0.04675 0.03986 -0.0513 0.9664 0.0316
-8.000 -0.6280 0.04385 0.03668 -0.0519 0.9627 0.0324
-7.750 -0.6161 0.04098 0.03341 -0.0507 0.9565 0.0325
-7.500 -0.5962 0.03817 0.03018 -0.0506 0.9528 0.0329
-7.250 -0.5718 0.03560 0.02717 -0.0511 0.9502 0.0335
-7.000 -0.5558 0.03383 0.02510 -0.0496 0.9443 0.0342
-6.750 -0.5318 0.03221 0.02317 -0.0495 0.9406 0.0362
-6.500 -0.5039 0.03048 0.02108 -0.0499 0.9379 0.0379
-6.250 -0.4730 0.02872 0.01900 -0.0507 0.9361 0.0388
-6.000 -0.4527 0.02754 0.01760 -0.0493 0.9313 0.0396
-5.750 -0.4266 0.02644 0.01630 -0.0491 0.9280 0.0405
-5.500 -0.3994 0.02497 0.01487 -0.0495 0.9256 0.0429
-5.250 -0.3717 0.02406 0.01395 -0.0499 0.9233 0.0462
-5.000 -0.3431 0.02315 0.01297 -0.0504 0.9213 0.0487
-4.750 -0.3293 0.02262 0.01237 -0.0480 0.9155 0.0506
-4.500 -0.3090 0.02201 0.01169 -0.0469 0.9117 0.0535
-4.250 -0.2861 0.02140 0.01107 -0.0465 0.9089 0.0591
-4.000 -0.2601 0.02091 0.01044 -0.0465 0.9068 0.0648
-3.750 -0.2461 0.02054 0.01009 -0.0442 0.9021 0.0717
-3.500 -0.2281 0.02009 0.00972 -0.0427 0.8979 0.0913
-3.000 -0.2113 0.01749 0.01005 -0.0352 0.8917 0.6785
-2.750 -0.2042 0.01814 0.01072 -0.0302 0.8867 0.7512
-2.500 -0.1899 0.01914 0.01173 -0.0254 0.8832 0.7991
-2.250 -0.1659 0.02005 0.01257 -0.0226 0.8809 0.8292
-2.000 -0.1380 0.02067 0.01309 -0.0211 0.8790 0.8512
-1.750 -0.1040 0.02093 0.01320 -0.0218 0.8777 0.8606
-1.500 -0.0848 0.02096 0.01313 -0.0205 0.8746 0.8668
-1.250 -0.0680 0.02108 0.01318 -0.0188 0.8702 0.8721
-1.000 -0.0506 0.02103 0.01305 -0.0172 0.8664 0.8788
-0.750 -0.0206 0.02106 0.01299 -0.0180 0.8644 0.8821
-0.500 0.0110 0.02107 0.01293 -0.0191 0.8627 0.8852
-0.250 0.0230 0.02114 0.01297 -0.0166 0.8578 0.8909
0.000 0.0381 0.02116 0.01296 -0.0148 0.8530 0.8957
0.250 0.0688 0.02117 0.01294 -0.0157 0.8504 0.8982
0.500 0.1006 0.02116 0.01291 -0.0169 0.8482 0.9007
0.750 0.1168 0.02123 0.01298 -0.0152 0.8434 0.9053
1.000 0.1316 0.02130 0.01304 -0.0133 0.8379 0.9103
1.250 0.1640 0.02132 0.01309 -0.0145 0.8351 0.9126
1.500 0.1980 0.02130 0.01310 -0.0160 0.8329 0.9149
1.750 0.2090 0.02146 0.01330 -0.0135 0.8258 0.9205
2.000 0.2334 0.02148 0.01336 -0.0132 0.8214 0.9245
2.250 0.2701 0.02145 0.01341 -0.0152 0.8188 0.9264
2.750 0.3168 0.02164 0.01377 -0.0144 0.8072 0.9351
3.000 0.3495 0.02152 0.01375 -0.0153 0.8038 0.9382
3.250 0.3747 0.02164 0.01401 -0.0153 0.7964 0.9418
3.500 0.4096 0.02147 0.01399 -0.0166 0.7907 0.9443
3.750 0.4426 0.02098 0.01365 -0.0170 0.7796 0.9475
4.000 0.4996 0.01889 0.01171 -0.0192 0.7570 0.9457
4.250 0.5187 0.01784 0.01074 -0.0160 0.7222 0.9516
4.500 0.5379 0.01700 0.00993 -0.0132 0.6497 0.9567
4.750 0.5626 0.01721 0.00848 -0.0105 0.3550 0.9585
5.000 0.5654 0.01905 0.00921 -0.0074 0.1515 0.9662
5.250 0.5853 0.02035 0.00999 -0.0070 0.0774 0.9715
5.500 0.6105 0.02117 0.01077 -0.0072 0.0632 0.9768
5.750 0.6367 0.02206 0.01171 -0.0078 0.0566 0.9821
6.000 0.6632 0.02288 0.01264 -0.0084 0.0504 0.9877
6.250 0.6864 0.02400 0.01377 -0.0087 0.0468 0.9948
6.500 0.7011 0.02476 0.01464 -0.0069 0.0449 1.0000
6.750 0.7040 0.02528 0.01522 -0.0028 0.0431 1.0000
7.000 0.7160 0.02603 0.01599 -0.0005 0.0407 1.0000
7.250 0.7342 0.02702 0.01699 0.0006 0.0381 1.0000
7.500 0.7657 0.02892 0.01886 -0.0006 0.0363 1.0000
7.750 0.7962 0.03057 0.02066 -0.0012 0.0354 1.0000
8.000 0.8249 0.03231 0.02263 -0.0015 0.0345 1.0000
8.250 0.8504 0.03419 0.02481 -0.0014 0.0335 1.0000
8.500 0.8714 0.03597 0.02686 -0.0007 0.0320 1.0000
8.750 0.8899 0.03781 0.02896 0.0002 0.0305 1.0000
9.000 0.9067 0.04001 0.03146 0.0013 0.0298 1.0000
9.250 0.9204 0.04241 0.03418 0.0028 0.0294 1.0000
9.500 0.9306 0.04498 0.03708 0.0046 0.0291 1.0000
9.750 0.9372 0.04770 0.04012 0.0067 0.0288 1.0000
10.000 0.9394 0.05057 0.04335 0.0092 0.0287 1.0000
10.250 0.9367 0.05351 0.04661 0.0121 0.0286 1.0000
10.500 0.9281 0.05644 0.04983 0.0154 0.0286 1.0000
10.750 0.9169 0.05946 0.05311 0.0185 0.0286 1.0000
11.000 0.9066 0.06238 0.05623 0.0208 0.0284 1.0000
11.250 0.8914 0.06578 0.05985 0.0228 0.0283 1.0000
11.500 0.8765 0.06930 0.06356 0.0241 0.0282 1.0000
11.750 0.8487 0.07439 0.06891 0.0246 0.0286 1.0000
12.000 0.8221 0.08005 0.07477 0.0236 0.0288 1.0000
12.250 0.7847 0.08860 0.08351 0.0198 0.0298 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 66(1)-212 (naca661212-il)