NACA 66-210 (naca66210-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 66-210 (naca66210-il) Reynolds number: 500,000 Max Cl/Cd: 61.92 at α=2.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-naca66210-il-500000-n5.txt Download as CSV file: xf-naca66210-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 66-210
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.5502 0.08425 0.08199 -0.0412 1.0000 0.0066
-10.250 -0.5456 0.08151 0.07926 -0.0431 1.0000 0.0072
-10.000 -0.5697 0.07132 0.06910 -0.0517 1.0000 0.0069
-9.750 -0.6286 0.06028 0.05796 -0.0557 1.0000 0.0064
-9.500 -0.6319 0.05703 0.05463 -0.0565 0.9865 0.0066
-9.250 -0.6240 0.05247 0.04989 -0.0597 0.9733 0.0068
-9.000 -0.6225 0.04613 0.04324 -0.0617 0.9604 0.0071
-8.750 -0.6235 0.03984 0.03655 -0.0611 0.9474 0.0075
-8.500 -0.6518 0.02667 0.02224 -0.0552 0.9309 0.0083
-8.250 -0.6420 0.02245 0.01736 -0.0530 0.9232 0.0089
-8.000 -0.6249 0.02046 0.01505 -0.0517 0.9164 0.0094
-7.750 -0.6023 0.02007 0.01459 -0.0513 0.9109 0.0098
-7.500 -0.5790 0.01976 0.01420 -0.0509 0.9058 0.0103
-7.250 -0.5558 0.01913 0.01344 -0.0504 0.9006 0.0110
-7.000 -0.5330 0.01800 0.01208 -0.0498 0.8962 0.0117
-6.750 -0.5096 0.01693 0.01081 -0.0492 0.8917 0.0124
-6.500 -0.4855 0.01608 0.00981 -0.0487 0.8871 0.0129
-6.250 -0.4625 0.01503 0.00861 -0.0481 0.8829 0.0136
-6.000 -0.4381 0.01453 0.00806 -0.0478 0.8788 0.0145
-5.750 -0.4131 0.01417 0.00764 -0.0475 0.8743 0.0154
-5.500 -0.3885 0.01363 0.00703 -0.0471 0.8702 0.0161
-5.250 -0.3641 0.01307 0.00639 -0.0467 0.8668 0.0168
-5.000 -0.3396 0.01255 0.00580 -0.0463 0.8629 0.0175
-4.750 -0.3145 0.01216 0.00536 -0.0459 0.8590 0.0184
-4.500 -0.2888 0.01188 0.00500 -0.0457 0.8556 0.0191
-4.250 -0.2658 0.01118 0.00423 -0.0450 0.8525 0.0204
-4.000 -0.2408 0.01077 0.00380 -0.0447 0.8488 0.0220
-3.750 -0.2151 0.01048 0.00348 -0.0445 0.8450 0.0236
-3.500 -0.1890 0.01023 0.00317 -0.0443 0.8417 0.0256
-3.250 -0.1626 0.01003 0.00291 -0.0442 0.8389 0.0274
-3.000 -0.1360 0.00983 0.00268 -0.0441 0.8360 0.0299
-2.750 -0.1096 0.00958 0.00244 -0.0440 0.8328 0.0353
-2.500 -0.0828 0.00942 0.00224 -0.0440 0.8298 0.0399
-2.250 -0.0563 0.00918 0.00208 -0.0439 0.8271 0.0603
-2.000 -0.0337 0.00839 0.00184 -0.0436 0.8244 0.2236
-1.750 -0.0143 0.00717 0.00156 -0.0429 0.8211 0.4874
-1.500 0.0062 0.00643 0.00157 -0.0418 0.8178 0.6956
-1.250 0.0318 0.00639 0.00168 -0.0412 0.8151 0.7521
-1.000 0.0587 0.00641 0.00175 -0.0409 0.8127 0.7730
-0.750 0.0866 0.00644 0.00176 -0.0409 0.8106 0.7836
-0.500 0.1144 0.00647 0.00178 -0.0410 0.8082 0.7927
-0.250 0.1416 0.00651 0.00185 -0.0408 0.8053 0.8038
0.000 0.1699 0.00653 0.00186 -0.0410 0.8026 0.8091
0.250 0.1982 0.00653 0.00188 -0.0412 0.8001 0.8115
0.500 0.2263 0.00654 0.00188 -0.0414 0.7969 0.8139
0.750 0.2540 0.00653 0.00188 -0.0414 0.7904 0.8164
1.000 0.2814 0.00652 0.00183 -0.0413 0.7798 0.8191
1.250 0.3088 0.00651 0.00182 -0.0412 0.7689 0.8220
1.500 0.3363 0.00651 0.00183 -0.0412 0.7587 0.8249
1.750 0.3631 0.00651 0.00184 -0.0410 0.7447 0.8274
2.000 0.3890 0.00654 0.00183 -0.0405 0.7190 0.8302
2.250 0.4130 0.00667 0.00179 -0.0397 0.6647 0.8334
2.500 0.4218 0.00791 0.00205 -0.0362 0.4429 0.8376
2.750 0.4339 0.00920 0.00251 -0.0339 0.2427 0.8411
3.000 0.4523 0.01002 0.00286 -0.0326 0.1218 0.8447
3.250 0.4733 0.01065 0.00320 -0.0316 0.0476 0.8488
3.500 0.4982 0.01094 0.00344 -0.0313 0.0355 0.8528
3.750 0.5231 0.01115 0.00369 -0.0309 0.0317 0.8559
4.000 0.5474 0.01143 0.00397 -0.0304 0.0273 0.8595
4.250 0.5715 0.01177 0.00438 -0.0298 0.0243 0.8636
4.500 0.5964 0.01202 0.00467 -0.0294 0.0227 0.8676
4.750 0.6201 0.01232 0.00503 -0.0288 0.0210 0.8713
5.000 0.6435 0.01265 0.00540 -0.0281 0.0194 0.8757
5.250 0.6665 0.01307 0.00586 -0.0274 0.0180 0.8808
5.500 0.6863 0.01372 0.00658 -0.0261 0.0168 0.8855
5.750 0.7066 0.01435 0.00729 -0.0248 0.0162 0.8909
6.000 0.7292 0.01478 0.00779 -0.0241 0.0157 0.8962
6.250 0.7506 0.01524 0.00833 -0.0230 0.0150 0.9014
6.500 0.7716 0.01584 0.00902 -0.0220 0.0145 0.9074
6.750 0.7922 0.01650 0.00978 -0.0209 0.0141 0.9132
7.000 0.8129 0.01717 0.01053 -0.0198 0.0136 0.9198
7.250 0.8339 0.01790 0.01136 -0.0188 0.0132 0.9267
7.500 0.8545 0.01848 0.01203 -0.0177 0.0127 0.9347
7.750 0.8752 0.01897 0.01258 -0.0167 0.0122 0.9435
8.000 0.8959 0.01968 0.01337 -0.0158 0.0116 0.9539
8.500 0.9464 0.02238 0.01645 -0.0161 0.0107 0.9781
8.750 0.9706 0.02382 0.01811 -0.0161 0.0102 1.0000
9.000 0.9912 0.02513 0.01961 -0.0154 0.0097 1.0000
9.250 1.0107 0.02628 0.02092 -0.0145 0.0091 1.0000
9.500 1.0293 0.02725 0.02203 -0.0136 0.0087 1.0000
9.750 1.0475 0.02781 0.02270 -0.0127 0.0083 1.0000
10.000 1.0640 0.02864 0.02362 -0.0116 0.0081 1.0000
10.250 1.0784 0.02967 0.02478 -0.0102 0.0078 1.0000
10.500 1.0812 0.03245 0.02786 -0.0075 0.0074 1.0000
10.750 1.0829 0.03512 0.03089 -0.0044 0.0073 1.0000
11.000 1.0719 0.03911 0.03531 -0.0002 0.0071 1.0000
11.250 1.0227 0.04812 0.04499 0.0066 0.0067 1.0000
11.500 0.9752 0.05632 0.05360 0.0107 0.0066 1.0000
11.750 0.9437 0.06242 0.05992 0.0118 0.0066 1.0000
12.000 0.9097 0.06967 0.06736 0.0107 0.0066 1.0000
12.250 0.8801 0.07748 0.07529 0.0072 0.0067 1.0000
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Polar data table (+)
Polar graphs
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