NACA 63-212 AIRFOIL (n63212-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 63-212 AIRFOIL (n63212-il) Reynolds number: 50,000 Max Cl/Cd: 32.81 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n63212-il-50000-n5.txt Download as CSV file: xf-n63212-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63-212 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.5934 0.09821 0.09080 -0.0286 1.0000 0.0484
-11.000 -0.6009 0.09205 0.08466 -0.0322 1.0000 0.0482
-10.750 -0.6104 0.08609 0.07870 -0.0359 1.0000 0.0478
-10.500 -0.6251 0.08005 0.07263 -0.0399 1.0000 0.0475
-10.250 -0.6420 0.07463 0.06718 -0.0432 1.0000 0.0472
-10.000 -0.6592 0.07000 0.06246 -0.0453 1.0000 0.0470
-9.750 -0.6765 0.06607 0.05842 -0.0459 1.0000 0.0468
-9.500 -0.6881 0.06237 0.05453 -0.0458 1.0000 0.0468
-9.250 -0.6962 0.05869 0.05059 -0.0455 1.0000 0.0470
-9.000 -0.6999 0.05514 0.04666 -0.0448 1.0000 0.0474
-8.750 -0.6990 0.05185 0.04297 -0.0439 1.0000 0.0479
-8.500 -0.6942 0.04872 0.03944 -0.0428 1.0000 0.0486
-8.250 -0.6825 0.04586 0.03643 -0.0421 1.0000 0.0496
-8.000 -0.6693 0.04330 0.03366 -0.0411 1.0000 0.0505
-7.750 -0.6544 0.04087 0.03097 -0.0401 1.0000 0.0513
-7.500 -0.6379 0.03864 0.02850 -0.0389 1.0000 0.0524
-7.250 -0.6202 0.03662 0.02626 -0.0376 1.0000 0.0537
-7.000 -0.6014 0.03478 0.02419 -0.0363 1.0000 0.0554
-6.750 -0.5821 0.03316 0.02236 -0.0347 1.0000 0.0575
-6.500 -0.5639 0.03175 0.02084 -0.0330 1.0000 0.0607
-6.250 -0.5486 0.03058 0.01972 -0.0312 1.0000 0.0649
-6.000 -0.5333 0.02954 0.01851 -0.0291 1.0000 0.0702
-5.750 -0.5215 0.02846 0.01748 -0.0269 1.0000 0.0764
-5.500 -0.5097 0.02754 0.01648 -0.0247 1.0000 0.0858
-5.250 -0.4992 0.02654 0.01551 -0.0226 1.0000 0.0977
-5.000 -0.4885 0.02545 0.01452 -0.0206 1.0000 0.1139
-4.750 -0.4775 0.02414 0.01340 -0.0189 1.0000 0.1393
-4.500 -0.4666 0.02232 0.01218 -0.0179 1.0000 0.2111
-4.250 -0.4428 0.02072 0.01204 -0.0184 0.9907 0.4527
-4.000 -0.4130 0.02100 0.01250 -0.0181 0.9806 0.5638
-3.750 -0.3849 0.02169 0.01321 -0.0168 0.9708 0.6304
-3.500 -0.3560 0.02243 0.01385 -0.0154 0.9618 0.6761
-3.250 -0.3278 0.02293 0.01422 -0.0142 0.9523 0.7042
-3.000 -0.2972 0.02312 0.01423 -0.0142 0.9434 0.7251
-2.750 -0.2619 0.02320 0.01408 -0.0154 0.9356 0.7411
-2.500 -0.2316 0.02317 0.01387 -0.0159 0.9261 0.7554
-2.000 -0.1653 0.02314 0.01351 -0.0179 0.9095 0.7797
-1.750 -0.1289 0.02309 0.01330 -0.0196 0.9026 0.7912
-1.500 -0.1021 0.02302 0.01313 -0.0197 0.8924 0.8032
-1.250 -0.0712 0.02298 0.01296 -0.0204 0.8842 0.8148
-1.000 -0.0403 0.02294 0.01285 -0.0210 0.8756 0.8253
-0.750 -0.0125 0.02291 0.01275 -0.0212 0.8665 0.8368
-0.500 0.0193 0.02286 0.01262 -0.0220 0.8587 0.8486
-0.250 0.0451 0.02286 0.01259 -0.0219 0.8490 0.8606
0.000 0.0805 0.02282 0.01251 -0.0232 0.8423 0.8712
0.250 0.1065 0.02285 0.01253 -0.0231 0.8321 0.8837
0.500 0.1426 0.02281 0.01246 -0.0247 0.8257 0.8954
0.750 0.1703 0.02288 0.01254 -0.0251 0.8155 0.9088
1.000 0.2084 0.02287 0.01254 -0.0271 0.8086 0.9205
1.250 0.2441 0.02293 0.01264 -0.0291 0.7994 0.9328
1.500 0.2847 0.02295 0.01271 -0.0318 0.7919 0.9443
1.750 0.3250 0.02301 0.01284 -0.0346 0.7830 0.9563
2.000 0.3672 0.02307 0.01300 -0.0379 0.7746 0.9680
2.250 0.4114 0.02310 0.01312 -0.0414 0.7662 0.9795
2.500 0.4519 0.02323 0.01337 -0.0446 0.7563 0.9933
2.750 0.4811 0.02326 0.01349 -0.0452 0.7482 1.0000
3.000 0.4887 0.02363 0.01393 -0.0425 0.7362 1.0000
3.250 0.5008 0.02395 0.01430 -0.0403 0.7253 1.0000
3.500 0.5244 0.02401 0.01443 -0.0394 0.7170 1.0000
3.750 0.5403 0.02440 0.01493 -0.0380 0.7046 1.0000
4.000 0.5607 0.02472 0.01534 -0.0370 0.6926 1.0000
4.250 0.5850 0.02489 0.01563 -0.0363 0.6809 1.0000
4.500 0.6118 0.02489 0.01576 -0.0356 0.6687 1.0000
4.750 0.6368 0.02487 0.01590 -0.0345 0.6533 1.0000
5.000 0.6600 0.02480 0.01597 -0.0331 0.6340 1.0000
5.250 0.6865 0.02431 0.01560 -0.0314 0.6115 1.0000
5.500 0.7095 0.02378 0.01519 -0.0291 0.5811 1.0000
5.750 0.7292 0.02337 0.01482 -0.0266 0.5405 1.0000
6.000 0.7480 0.02310 0.01450 -0.0239 0.4874 1.0000
6.250 0.7634 0.02327 0.01448 -0.0213 0.4108 1.0000
6.500 0.7724 0.02425 0.01474 -0.0183 0.2906 1.0000
6.750 0.7748 0.02635 0.01602 -0.0158 0.1946 1.0000
7.000 0.7798 0.02846 0.01767 -0.0138 0.1464 1.0000
7.250 0.7883 0.03030 0.01929 -0.0121 0.1203 1.0000
7.500 0.7989 0.03200 0.02089 -0.0106 0.1050 1.0000
7.750 0.8120 0.03359 0.02249 -0.0092 0.0938 1.0000
8.000 0.8266 0.03515 0.02403 -0.0079 0.0849 1.0000
8.250 0.8468 0.03665 0.02561 -0.0070 0.0779 1.0000
8.500 0.8721 0.03825 0.02728 -0.0066 0.0715 1.0000
8.750 0.8983 0.03989 0.02909 -0.0063 0.0660 1.0000
9.000 0.9292 0.04202 0.03116 -0.0065 0.0624 1.0000
9.250 0.9532 0.04420 0.03375 -0.0061 0.0589 1.0000
9.500 0.9729 0.04644 0.03626 -0.0056 0.0557 1.0000
9.750 0.9914 0.04892 0.03893 -0.0051 0.0536 1.0000
10.000 1.0086 0.05174 0.04196 -0.0046 0.0523 1.0000
10.250 1.0223 0.05506 0.04547 -0.0041 0.0513 1.0000
10.500 1.0237 0.05829 0.04916 -0.0026 0.0506 1.0000
10.750 1.0183 0.06157 0.05285 -0.0008 0.0500 1.0000
11.000 1.0083 0.06495 0.05656 0.0008 0.0495 1.0000
11.250 0.9945 0.06867 0.06058 0.0017 0.0491 1.0000
11.500 0.9777 0.07287 0.06505 0.0018 0.0489 1.0000
11.750 0.9584 0.07770 0.07012 0.0008 0.0489 1.0000
12.000 0.9380 0.08318 0.07580 -0.0012 0.0491 1.0000
12.250 0.9147 0.08964 0.08244 -0.0044 0.0493 1.0000
12.500 0.8918 0.09688 0.08982 -0.0087 0.0497 1.0000
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Polar data table (+)
Polar graphs
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