NACA 6-H-10 AIRFOIL (n6h10-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 6-H-10 AIRFOIL (n6h10-il) Reynolds number: 500,000 Max Cl/Cd: 87 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n6h10-il-500000.txt Download as CSV file: xf-n6h10-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 6-H-10 AIRFOIL
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
------ -------- --------- --------- -------- -------- --------
-7.750 -0.4575 0.09119 0.08808 0.0098 0.6436 0.0109
-7.500 -0.4602 0.08809 0.08499 0.0086 0.6415 0.0110
-7.250 -0.4620 0.08508 0.08196 0.0080 0.6396 0.0111
-7.000 -0.4580 0.08172 0.07856 0.0066 0.6378 0.0114
-6.750 -0.4519 0.07830 0.07509 0.0052 0.6361 0.0116
-6.500 -0.4429 0.07490 0.07162 0.0038 0.6344 0.0119
-6.250 -0.4301 0.07141 0.06806 0.0022 0.6326 0.0122
-6.000 -0.4148 0.06796 0.06452 0.0009 0.6306 0.0124
-5.750 -0.3984 0.06448 0.06094 0.0001 0.6287 0.0125
-5.500 -0.3819 0.06109 0.05745 -0.0002 0.6267 0.0125
-5.250 -0.3648 0.05766 0.05388 -0.0001 0.6248 0.0126
-5.000 -0.3471 0.05430 0.05037 0.0002 0.6231 0.0126
-4.750 -0.3355 0.04927 0.04518 0.0009 0.6215 0.0128
-4.500 -0.3230 0.04601 0.04182 0.0015 0.6198 0.0132
-4.250 -0.3052 0.04348 0.03921 0.0019 0.6180 0.0136
-4.000 -0.2857 0.04097 0.03659 0.0025 0.6161 0.0142
-3.750 -0.2646 0.03839 0.03385 0.0032 0.6142 0.0149
-3.500 -0.2415 0.03573 0.03098 0.0044 0.6125 0.0164
-3.250 -0.2098 0.03413 0.02907 0.0061 0.6108 0.0180
-3.000 -0.1944 0.02907 0.02363 0.0083 0.6094 0.0188
-2.750 -0.1735 0.02737 0.02184 0.0087 0.6079 0.0198
-2.500 -0.1502 0.02612 0.02045 0.0092 0.6062 0.0217
-2.250 -0.1220 0.02495 0.01907 0.0102 0.6044 0.0250
-2.000 -0.0934 0.02478 0.01863 0.0111 0.6023 0.0260
-1.750 -0.0731 0.02057 0.01417 0.0121 0.6005 0.0285
-1.500 -0.0470 0.01946 0.01296 0.0124 0.5986 0.0309
-1.250 -0.0189 0.01898 0.01229 0.0129 0.5968 0.0356
-1.000 0.0075 0.01710 0.01019 0.0133 0.5952 0.0410
-0.750 0.0424 0.01449 0.00716 0.0142 0.5936 0.0229
-0.500 0.0716 0.01349 0.00602 0.0143 0.5919 0.0225
-0.250 0.0998 0.01301 0.00554 0.0143 0.5900 0.0245
0.000 0.1266 0.01229 0.00480 0.0147 0.5882 0.0244
0.250 0.1525 0.01182 0.00430 0.0152 0.5866 0.0252
0.500 0.1777 0.01143 0.00389 0.0158 0.5851 0.0267
0.750 0.2028 0.01110 0.00354 0.0164 0.5836 0.0322
1.000 0.2290 0.01093 0.00333 0.0168 0.5822 0.0397
1.250 0.2298 0.00906 0.00310 0.0217 0.5810 0.6114
1.500 0.2346 0.00863 0.00354 0.0278 0.5797 0.9090
1.750 0.2644 0.00924 0.00415 0.0286 0.5782 0.9398
2.000 0.3980 0.01076 0.00559 0.0076 0.5764 0.9597
2.250 0.4415 0.01092 0.00574 0.0041 0.5747 0.9630
2.500 0.4717 0.01098 0.00581 0.0034 0.5731 0.9650
2.750 0.4983 0.01106 0.00590 0.0035 0.5714 0.9676
3.000 0.5192 0.01111 0.00595 0.0048 0.5697 0.9693
3.250 0.5416 0.01110 0.00593 0.0057 0.5679 0.9700
3.500 0.5658 0.01100 0.00577 0.0064 0.5650 0.9700
3.750 0.5908 0.01080 0.00563 0.0068 0.5588 0.9701
4.000 0.6160 0.01057 0.00537 0.0072 0.5530 0.9702
4.250 0.6424 0.01047 0.00525 0.0073 0.5485 0.9704
4.500 0.6700 0.01035 0.00520 0.0071 0.5434 0.9707
4.750 0.6968 0.01021 0.00506 0.0072 0.5380 0.9710
5.000 0.7232 0.01010 0.00502 0.0072 0.5311 0.9712
5.250 0.7497 0.00996 0.00486 0.0073 0.5227 0.9715
5.500 0.7768 0.00987 0.00487 0.0071 0.5139 0.9718
5.750 0.8039 0.00980 0.00484 0.0070 0.5027 0.9722
6.000 0.8312 0.00975 0.00485 0.0068 0.4863 0.9726
6.250 0.8578 0.00986 0.00486 0.0066 0.4392 0.9730
6.500 0.8716 0.01304 0.00677 0.0052 0.1892 0.9742
6.750 0.8869 0.01510 0.00826 0.0046 0.0798 0.9757
7.000 0.9071 0.01596 0.00912 0.0046 0.0691 0.9771
7.250 0.9284 0.01648 0.00972 0.0048 0.0657 0.9785
7.500 0.9461 0.01715 0.01044 0.0054 0.0622 0.9805
7.750 0.9560 0.01798 0.01132 0.0072 0.0593 0.9834
8.000 0.9798 0.01961 0.01305 0.0045 0.0559 0.9849
8.250 1.0049 0.02065 0.01420 0.0025 0.0543 0.9865
8.500 1.0265 0.02160 0.01522 0.0013 0.0517 0.9885
8.750 1.0419 0.02256 0.01623 0.0011 0.0493 0.9914
9.000 1.0524 0.02385 0.01755 0.0013 0.0469 0.9946
9.250 1.0620 0.02595 0.01972 0.0004 0.0442 0.9974
9.500 1.0867 0.02655 0.02040 -0.0008 0.0429 0.9988
9.750 1.1115 0.02720 0.02113 -0.0021 0.0402 1.0000
10.000 1.0976 0.02801 0.02199 0.0034 0.0396 1.0000
10.250 1.0857 0.02858 0.02257 0.0088 0.0383 1.0000
10.500 1.0865 0.02940 0.02338 0.0117 0.0362 1.0000
10.750 1.0816 0.03101 0.02501 0.0146 0.0338 1.0000
11.000 1.0991 0.03139 0.02546 0.0152 0.0309 1.0000
11.250 1.1124 0.03211 0.02614 0.0160 0.0260 1.0000
11.500 1.1219 0.03314 0.02719 0.0172 0.0207 1.0000
11.750 1.1256 0.03467 0.02870 0.0185 0.0180 1.0000
12.000 1.1315 0.03608 0.03016 0.0197 0.0158 1.0000
12.250 1.1365 0.03763 0.03171 0.0208 0.0138 1.0000
12.500 1.1337 0.03988 0.03402 0.0222 0.0125 1.0000
12.750 1.1360 0.04180 0.03604 0.0233 0.0120 1.0000
13.000 1.1417 0.04348 0.03782 0.0242 0.0114 1.0000
13.250 1.1464 0.04530 0.03973 0.0250 0.0107 1.0000
13.500 1.1511 0.04716 0.04168 0.0258 0.0104 1.0000
13.750 1.1546 0.04917 0.04379 0.0266 0.0101 1.0000
14.000 1.1571 0.05137 0.04604 0.0271 0.0092 1.0000
14.250 1.1616 0.05337 0.04814 0.0276 0.0091 1.0000
14.500 1.1554 0.05643 0.05136 0.0290 0.0085 1.0000
14.750 1.1544 0.05917 0.05429 0.0298 0.0083 1.0000
15.000 1.1586 0.06150 0.05676 0.0297 0.0079 1.0000
15.250 1.1579 0.06439 0.05980 0.0299 0.0079 1.0000
15.500 1.1571 0.06741 0.06299 0.0298 0.0075 1.0000
15.750 1.1509 0.07118 0.06695 0.0298 0.0077 1.0000
16.000 1.1477 0.07473 0.07064 0.0292 0.0073 1.0000
16.250 1.1405 0.07891 0.07500 0.0284 0.0072 1.0000
16.500 1.1320 0.08346 0.07971 0.0273 0.0071 1.0000
16.750 1.1204 0.08865 0.08509 0.0259 0.0070 1.0000
17.000 1.1011 0.09533 0.09200 0.0239 0.0071 1.0000
17.250 1.0895 0.10102 0.09783 0.0216 0.0070 1.0000
17.500 1.0768 0.10714 0.10411 0.0190 0.0070 1.0000
17.750 1.0496 0.11645 0.11365 0.0149 0.0072 1.0000
18.000 1.0386 0.12289 0.12019 0.0115 0.0069 1.0000
18.250 1.0123 0.13313 0.13062 0.0063 0.0071 1.0000
18.500 0.9871 0.14395 0.14160 0.0005 0.0072 1.0000
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