NACA 65(1)-212 a=0.6 (naca651212a06-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 65(1)-212 a=0.6 (naca651212a06-il) Reynolds number: 200,000 Max Cl/Cd: 56.65 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-naca651212a06-il-200000-n5.txt Download as CSV file: xf-naca651212a06-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 65(1)-212 a=0.6
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.5422 0.08996 0.08643 -0.0349 1.0000 0.0188
-10.750 -0.5552 0.08122 0.07774 -0.0408 1.0000 0.0187
-10.500 -0.5776 0.07089 0.06738 -0.0491 1.0000 0.0184
-10.250 -0.6001 0.06419 0.06058 -0.0537 1.0000 0.0181
-10.000 -0.6225 0.05902 0.05530 -0.0559 1.0000 0.0179
-9.750 -0.6448 0.05482 0.05095 -0.0553 1.0000 0.0179
-9.500 -0.6558 0.05028 0.04616 -0.0554 0.9729 0.0179
-9.250 -0.6555 0.04506 0.04050 -0.0566 0.9520 0.0181
-9.000 -0.6539 0.04075 0.03575 -0.0557 0.9363 0.0184
-8.750 -0.6504 0.03711 0.03163 -0.0539 0.9228 0.0188
-8.500 -0.6429 0.03410 0.02809 -0.0518 0.9115 0.0197
-8.250 -0.6320 0.03165 0.02516 -0.0500 0.9024 0.0205
-8.000 -0.6160 0.02966 0.02296 -0.0489 0.8940 0.0210
-7.750 -0.5980 0.02806 0.02113 -0.0479 0.8869 0.0215
-7.500 -0.5784 0.02661 0.01944 -0.0469 0.8794 0.0221
-7.250 -0.5579 0.02524 0.01783 -0.0459 0.8736 0.0226
-7.000 -0.5354 0.02397 0.01636 -0.0453 0.8669 0.0234
-6.750 -0.5128 0.02297 0.01516 -0.0445 0.8611 0.0246
-6.500 -0.4892 0.02199 0.01397 -0.0439 0.8555 0.0258
-6.250 -0.4650 0.02096 0.01275 -0.0433 0.8496 0.0264
-6.000 -0.4410 0.02009 0.01173 -0.0426 0.8448 0.0269
-5.750 -0.4184 0.01871 0.01031 -0.0420 0.8396 0.0280
-5.500 -0.3955 0.01788 0.00945 -0.0414 0.8339 0.0290
-5.250 -0.3724 0.01727 0.00877 -0.0407 0.8293 0.0307
-5.000 -0.3486 0.01673 0.00818 -0.0402 0.8245 0.0328
-4.750 -0.3249 0.01617 0.00755 -0.0396 0.8193 0.0344
-4.500 -0.3013 0.01566 0.00694 -0.0389 0.8149 0.0357
-4.250 -0.2787 0.01502 0.00623 -0.0382 0.8105 0.0380
-4.000 -0.2543 0.01458 0.00576 -0.0378 0.8053 0.0415
-3.750 -0.2292 0.01425 0.00535 -0.0374 0.8010 0.0464
-3.500 -0.2046 0.01384 0.00491 -0.0369 0.7975 0.0540
-3.250 -0.1797 0.01343 0.00456 -0.0366 0.7929 0.0740
-3.000 -0.1599 0.01237 0.00413 -0.0359 0.7881 0.2173
-2.750 -0.1447 0.01100 0.00381 -0.0343 0.7840 0.4687
-2.500 -0.1225 0.01069 0.00386 -0.0332 0.7805 0.5863
-2.250 -0.0992 0.01060 0.00401 -0.0320 0.7760 0.6571
-2.000 -0.0750 0.01063 0.00413 -0.0310 0.7719 0.7003
-1.750 -0.0495 0.01068 0.00418 -0.0302 0.7683 0.7237
-1.500 -0.0227 0.01069 0.00414 -0.0299 0.7649 0.7361
-1.250 0.0051 0.01070 0.00408 -0.0300 0.7606 0.7436
-1.000 0.0327 0.01068 0.00403 -0.0300 0.7568 0.7497
-0.750 0.0605 0.01067 0.00395 -0.0300 0.7533 0.7569
-0.500 0.0883 0.01065 0.00387 -0.0301 0.7502 0.7634
-0.250 0.1159 0.01066 0.00386 -0.0302 0.7458 0.7710
0.000 0.1435 0.01064 0.00383 -0.0303 0.7419 0.7777
0.250 0.1713 0.01062 0.00378 -0.0303 0.7384 0.7855
0.500 0.1991 0.01057 0.00373 -0.0303 0.7355 0.7921
0.750 0.2266 0.01056 0.00377 -0.0304 0.7311 0.8000
1.000 0.2539 0.01053 0.00382 -0.0304 0.7270 0.8072
1.250 0.2813 0.01052 0.00386 -0.0302 0.7234 0.8148
1.500 0.3086 0.01054 0.00395 -0.0300 0.7205 0.8229
1.750 0.3349 0.01063 0.00419 -0.0295 0.7161 0.8312
2.000 0.3616 0.01075 0.00442 -0.0291 0.7119 0.8400
2.250 0.3893 0.01085 0.00462 -0.0287 0.7082 0.8499
2.750 0.4360 0.01116 0.00497 -0.0273 0.7000 0.9066
3.000 0.4620 0.01119 0.00507 -0.0269 0.6948 0.9160
3.250 0.4895 0.01115 0.00506 -0.0267 0.6897 0.9241
3.500 0.5158 0.01110 0.00511 -0.0263 0.6787 0.9325
3.750 0.5451 0.01086 0.00486 -0.0261 0.6583 0.9396
4.000 0.5730 0.01066 0.00463 -0.0258 0.6189 0.9475
4.250 0.6029 0.01066 0.00451 -0.0261 0.5652 0.9543
4.500 0.6283 0.01109 0.00451 -0.0258 0.4635 0.9617
4.750 0.6429 0.01280 0.00522 -0.0248 0.2712 0.9717
5.000 0.6608 0.01458 0.00618 -0.0248 0.1127 0.9817
5.250 0.6852 0.01565 0.00693 -0.0253 0.0570 0.9913
5.500 0.7104 0.01635 0.00759 -0.0255 0.0441 1.0000
5.750 0.7238 0.01691 0.00815 -0.0232 0.0387 1.0000
6.000 0.7385 0.01742 0.00872 -0.0211 0.0357 1.0000
6.250 0.7526 0.01803 0.00936 -0.0190 0.0331 1.0000
6.500 0.7638 0.01886 0.01024 -0.0166 0.0307 1.0000
6.750 0.7787 0.01950 0.01093 -0.0147 0.0292 1.0000
7.000 0.7940 0.02014 0.01163 -0.0130 0.0273 1.0000
7.250 0.8076 0.02089 0.01242 -0.0110 0.0261 1.0000
7.500 0.8205 0.02168 0.01326 -0.0090 0.0252 1.0000
7.750 0.8339 0.02255 0.01417 -0.0071 0.0243 1.0000
8.000 0.8482 0.02357 0.01520 -0.0054 0.0235 1.0000
8.250 0.8650 0.02491 0.01656 -0.0042 0.0228 1.0000
8.500 0.8858 0.02611 0.01784 -0.0034 0.0222 1.0000
8.750 0.9082 0.02729 0.01913 -0.0029 0.0218 1.0000
9.000 0.9302 0.02847 0.02047 -0.0023 0.0212 1.0000
9.250 0.9509 0.02970 0.02185 -0.0017 0.0203 1.0000
9.500 0.9704 0.03098 0.02328 -0.0009 0.0196 1.0000
9.750 0.9887 0.03240 0.02486 0.0000 0.0189 1.0000
10.000 1.0064 0.03405 0.02669 0.0009 0.0186 1.0000
10.250 1.0211 0.03574 0.02857 0.0021 0.0182 1.0000
10.500 1.0328 0.03745 0.03048 0.0035 0.0179 1.0000
10.750 1.0424 0.03934 0.03255 0.0051 0.0176 1.0000
11.000 1.0495 0.04137 0.03475 0.0067 0.0173 1.0000
11.250 1.0534 0.04373 0.03734 0.0085 0.0172 1.0000
11.500 1.0534 0.04659 0.04040 0.0103 0.0169 1.0000
11.750 1.0478 0.04952 0.04358 0.0123 0.0168 1.0000
12.000 1.0382 0.05227 0.04663 0.0143 0.0166 1.0000
12.250 1.0254 0.05556 0.05023 0.0159 0.0164 1.0000
12.500 1.0099 0.05930 0.05427 0.0169 0.0161 1.0000
12.750 0.9938 0.06353 0.05873 0.0173 0.0161 1.0000
13.000 0.9736 0.06844 0.06390 0.0168 0.0159 1.0000
13.250 0.9531 0.07381 0.06950 0.0155 0.0159 1.0000
13.500 0.9315 0.07985 0.07573 0.0132 0.0159 1.0000
13.750 0.9063 0.08718 0.08325 0.0096 0.0160 1.0000
14.000 0.8807 0.09572 0.09195 0.0046 0.0161 1.0000
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Polar data table (+)
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