NACA 5-H-10 AIRFOIL (n5h10-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 5-H-10 AIRFOIL (n5h10-il) Reynolds number: 200,000 Max Cl/Cd: 56.19 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n5h10-il-200000-n5.txt Download as CSV file: xf-n5h10-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 5-H-10 AIRFOIL
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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4955 0.09001 0.08648 -0.0160 0.8476 0.0147
-9.000 -0.4990 0.08556 0.08204 -0.0186 0.8419 0.0143
-8.750 -0.5073 0.08078 0.07726 -0.0218 0.8363 0.0142
-8.500 -0.5169 0.07639 0.07286 -0.0242 0.8297 0.0139
-8.250 -0.5289 0.07226 0.06868 -0.0247 0.8242 0.0137
-8.000 -0.5351 0.06753 0.06387 -0.0257 0.8187 0.0136
-7.750 -0.5386 0.06261 0.05882 -0.0262 0.8137 0.0134
-7.250 -0.5373 0.05254 0.04834 -0.0254 0.8046 0.0133
-7.000 -0.5323 0.04775 0.04327 -0.0242 0.7999 0.0135
-6.750 -0.5249 0.04330 0.03848 -0.0225 0.7959 0.0138
-6.500 -0.5135 0.03917 0.03401 -0.0209 0.7910 0.0149
-6.250 -0.5028 0.03425 0.02857 -0.0185 0.7864 0.0158
-6.000 -0.4904 0.02933 0.02295 -0.0158 0.7826 0.0165
-5.750 -0.4717 0.02607 0.01906 -0.0140 0.7785 0.0173
-5.500 -0.4511 0.02391 0.01658 -0.0130 0.7737 0.0185
-5.250 -0.4275 0.02355 0.01614 -0.0126 0.7693 0.0202
-5.000 -0.4032 0.02232 0.01463 -0.0119 0.7652 0.0224
-4.750 -0.3775 0.02072 0.01269 -0.0112 0.7600 0.0245
-4.500 -0.3513 0.02018 0.01187 -0.0107 0.7556 0.0267
-4.250 -0.3265 0.01858 0.01013 -0.0103 0.7523 0.0298
-4.000 -0.3000 0.01782 0.00928 -0.0101 0.7471 0.0326
-3.750 -0.2737 0.01726 0.00859 -0.0097 0.7425 0.0362
-3.500 -0.2475 0.01666 0.00782 -0.0093 0.7387 0.0379
-3.250 -0.2225 0.01572 0.00688 -0.0089 0.7338 0.0415
-3.000 -0.1971 0.01523 0.00636 -0.0085 0.7291 0.0443
-2.750 -0.1720 0.01475 0.00581 -0.0079 0.7253 0.0463
-2.500 -0.1467 0.01435 0.00533 -0.0074 0.7214 0.0482
-2.250 -0.1210 0.01405 0.00497 -0.0071 0.7165 0.0504
-2.000 -0.0961 0.01366 0.00456 -0.0065 0.7124 0.0544
-1.750 -0.0707 0.01338 0.00421 -0.0060 0.7091 0.0582
-1.500 -0.0446 0.01316 0.00395 -0.0057 0.7041 0.0645
-1.250 -0.0196 0.01282 0.00369 -0.0052 0.7000 0.0874
-1.000 -0.0002 0.01194 0.00346 -0.0040 0.6967 0.2683
-0.750 0.0031 0.01025 0.00389 0.0019 0.6935 0.8453
-0.500 0.0248 0.01046 0.00407 0.0038 0.6892 0.8882
-0.250 0.0597 0.01073 0.00429 0.0033 0.6855 0.9142
0.000 0.1397 0.01145 0.00489 -0.0061 0.6827 0.9463
0.250 0.2044 0.01174 0.00509 -0.0133 0.6786 0.9618
0.500 0.2357 0.01172 0.00501 -0.0142 0.6733 0.9637
0.750 0.2656 0.01167 0.00491 -0.0148 0.6690 0.9650
1.000 0.2920 0.01163 0.00482 -0.0147 0.6653 0.9656
1.250 0.3181 0.01161 0.00481 -0.0146 0.6605 0.9661
1.500 0.3446 0.01158 0.00476 -0.0144 0.6558 0.9668
1.750 0.3710 0.01156 0.00470 -0.0143 0.6523 0.9675
2.000 0.3971 0.01157 0.00476 -0.0142 0.6479 0.9682
2.250 0.4227 0.01157 0.00477 -0.0139 0.6428 0.9688
2.500 0.4482 0.01155 0.00473 -0.0136 0.6380 0.9695
2.750 0.4736 0.01154 0.00479 -0.0133 0.6301 0.9702
3.000 0.4986 0.01151 0.00474 -0.0127 0.6217 0.9710
3.250 0.5230 0.01146 0.00468 -0.0121 0.6079 0.9719
3.500 0.5472 0.01142 0.00467 -0.0114 0.5895 0.9730
3.750 0.5708 0.01140 0.00464 -0.0106 0.5685 0.9744
4.000 0.5961 0.01142 0.00467 -0.0103 0.5476 0.9754
4.250 0.6229 0.01146 0.00476 -0.0103 0.5262 0.9761
4.500 0.6490 0.01155 0.00481 -0.0101 0.4900 0.9769
4.750 0.6715 0.01199 0.00488 -0.0096 0.3956 0.9781
5.000 0.6912 0.01293 0.00538 -0.0090 0.3105 0.9796
5.250 0.7093 0.01408 0.00605 -0.0084 0.1926 0.9815
5.500 0.7282 0.01506 0.00673 -0.0078 0.1348 0.9835
5.750 0.7492 0.01566 0.00726 -0.0072 0.1040 0.9854
6.000 0.7727 0.01635 0.00785 -0.0073 0.0716 0.9869
6.250 0.7966 0.01708 0.00848 -0.0074 0.0506 0.9887
6.500 0.8198 0.01781 0.00922 -0.0075 0.0434 0.9909
6.750 0.8424 0.01846 0.00999 -0.0073 0.0396 0.9932
7.000 0.8637 0.01930 0.01090 -0.0071 0.0354 0.9956
7.250 0.8841 0.02037 0.01207 -0.0069 0.0318 0.9984
7.500 0.9020 0.02113 0.01296 -0.0060 0.0290 1.0000
7.750 0.9076 0.02179 0.01369 -0.0027 0.0263 1.0000
8.000 0.9082 0.02260 0.01458 0.0014 0.0245 1.0000
8.250 0.9039 0.02385 0.01584 0.0061 0.0229 1.0000
8.500 0.9123 0.02449 0.01658 0.0089 0.0217 1.0000
8.750 0.9194 0.02521 0.01740 0.0118 0.0200 1.0000
9.000 0.9264 0.02610 0.01837 0.0146 0.0182 1.0000
9.250 0.9352 0.02715 0.01949 0.0168 0.0171 1.0000
9.500 0.9451 0.02829 0.02069 0.0187 0.0161 1.0000
9.750 0.9549 0.02963 0.02208 0.0204 0.0151 1.0000
10.000 0.9671 0.03120 0.02377 0.0220 0.0140 1.0000
10.250 0.9800 0.03248 0.02522 0.0234 0.0131 1.0000
10.500 0.9924 0.03413 0.02710 0.0248 0.0124 1.0000
10.750 1.0028 0.03586 0.02903 0.0262 0.0118 1.0000
11.000 1.0109 0.03776 0.03112 0.0276 0.0113 1.0000
11.250 1.0167 0.03975 0.03330 0.0289 0.0109 1.0000
11.500 1.0204 0.04183 0.03556 0.0302 0.0106 1.0000
11.750 1.0221 0.04393 0.03781 0.0313 0.0103 1.0000
12.000 1.0210 0.04655 0.04063 0.0323 0.0101 1.0000
12.250 1.0163 0.04959 0.04386 0.0332 0.0099 1.0000
12.500 1.0074 0.05299 0.04749 0.0337 0.0096 1.0000
12.750 0.9954 0.05702 0.05182 0.0339 0.0094 1.0000
13.000 0.9829 0.06123 0.05626 0.0336 0.0094 1.0000
13.250 0.9680 0.06594 0.06122 0.0327 0.0093 1.0000
13.500 0.9493 0.07150 0.06704 0.0309 0.0089 1.0000
13.750 0.9333 0.07700 0.07268 0.0287 0.0093 1.0000
14.000 0.9126 0.08386 0.07973 0.0254 0.0093 1.0000
14.250 0.8921 0.09160 0.08765 0.0210 0.0093 1.0000
14.500 0.8738 0.09966 0.09583 0.0163 0.0095 1.0000
14.750 0.8532 0.10917 0.10543 0.0106 0.0097 1.0000
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Polar data table (+)
Polar graphs
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