NACA 5-H-10 AIRFOIL (n5h10-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 5-H-10 AIRFOIL (n5h10-il) Reynolds number: 1,000,000 Max Cl/Cd: 74.94 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n5h10-il-1000000.txt Download as CSV file: xf-n5h10-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.5180 0.08565 0.08349 -0.0150 0.7980 0.0114
-9.000 -0.5222 0.08160 0.07946 -0.0177 0.7943 0.0121
-7.750 -0.7101 0.02238 0.01805 -0.0076 0.7761 0.0093
-7.500 -0.6925 0.02014 0.01544 -0.0061 0.7728 0.0097
-7.250 -0.6742 0.01774 0.01263 -0.0046 0.7692 0.0103
-7.000 -0.6524 0.01631 0.01099 -0.0037 0.7653 0.0110
-6.750 -0.6271 0.01592 0.01054 -0.0034 0.7615 0.0114
-6.500 -0.6019 0.01552 0.01005 -0.0030 0.7578 0.0119
-6.250 -0.5766 0.01486 0.00927 -0.0026 0.7539 0.0126
-6.000 -0.5508 0.01433 0.00864 -0.0022 0.7496 0.0133
-5.750 -0.5250 0.01386 0.00806 -0.0018 0.7454 0.0137
-5.500 -0.5014 0.01262 0.00667 -0.0011 0.7411 0.0150
-5.250 -0.4742 0.01264 0.00671 -0.0010 0.7361 0.0159
-5.000 -0.4474 0.01252 0.00655 -0.0009 0.7314 0.0169
-4.750 -0.4208 0.01224 0.00622 -0.0006 0.7267 0.0180
-4.500 -0.3935 0.01215 0.00608 -0.0005 0.7213 0.0187
-4.250 -0.3698 0.01108 0.00492 0.0002 0.7164 0.0205
-4.000 -0.3429 0.01091 0.00475 0.0004 0.7110 0.0219
-3.750 -0.3159 0.01076 0.00458 0.0005 0.7052 0.0235
-3.500 -0.2891 0.01060 0.00436 0.0007 0.6999 0.0248
-3.250 -0.2613 0.01056 0.00431 0.0007 0.6941 0.0256
-3.000 -0.2388 0.00962 0.00326 0.0017 0.6888 0.0281
-2.750 -0.2125 0.00938 0.00301 0.0020 0.6838 0.0301
-2.500 -0.1860 0.00916 0.00276 0.0022 0.6785 0.0318
-2.250 -0.1597 0.00896 0.00251 0.0025 0.6731 0.0331
-2.000 -0.1328 0.00880 0.00233 0.0027 0.6674 0.0343
-1.750 -0.1060 0.00866 0.00215 0.0029 0.6617 0.0350
-1.500 -0.0802 0.00839 0.00180 0.0033 0.6566 0.0375
-1.250 -0.0535 0.00821 0.00162 0.0036 0.6515 0.0413
-1.000 -0.0266 0.00809 0.00148 0.0038 0.6468 0.0446
-0.750 0.0001 0.00797 0.00137 0.0040 0.6428 0.0541
-0.500 0.0233 0.00745 0.00126 0.0047 0.6391 0.1781
-0.250 0.0389 0.00641 0.00114 0.0067 0.6351 0.4675
0.000 0.0417 0.00516 0.00111 0.0120 0.6313 0.8435
0.250 0.0675 0.00526 0.00122 0.0126 0.6273 0.8670
0.500 0.0938 0.00533 0.00131 0.0131 0.6233 0.8803
0.750 0.1200 0.00541 0.00140 0.0137 0.6196 0.8915
1.000 0.1455 0.00550 0.00147 0.0144 0.6157 0.9018
1.250 0.1716 0.00555 0.00155 0.0150 0.6108 0.9107
1.500 0.1987 0.00557 0.00154 0.0152 0.6039 0.9139
1.750 0.2259 0.00557 0.00155 0.0154 0.5946 0.9184
2.000 0.2515 0.00565 0.00160 0.0160 0.5877 0.9252
2.250 0.2783 0.00564 0.00160 0.0163 0.5806 0.9292
2.500 0.3060 0.00567 0.00160 0.0163 0.5717 0.9305
2.750 0.3342 0.00568 0.00162 0.0161 0.5634 0.9317
3.000 0.3620 0.00572 0.00164 0.0160 0.5550 0.9331
3.250 0.3896 0.00576 0.00166 0.0160 0.5435 0.9344
3.500 0.4173 0.00582 0.00169 0.0159 0.5234 0.9351
3.750 0.4444 0.00593 0.00173 0.0158 0.4979 0.9357
4.000 0.4692 0.00627 0.00184 0.0161 0.4318 0.9364
4.250 0.4885 0.00718 0.00222 0.0170 0.3109 0.9376
4.500 0.5123 0.00763 0.00246 0.0173 0.2557 0.9384
4.750 0.5304 0.00863 0.00293 0.0183 0.1311 0.9398
5.000 0.5533 0.00913 0.00321 0.0188 0.0822 0.9407
5.500 0.6017 0.00991 0.00376 0.0194 0.0356 0.9430
5.750 0.6273 0.01017 0.00407 0.0195 0.0327 0.9442
6.000 0.6530 0.01040 0.00435 0.0196 0.0317 0.9454
6.250 0.6781 0.01067 0.00466 0.0198 0.0300 0.9469
6.500 0.7026 0.01099 0.00503 0.0200 0.0281 0.9485
6.750 0.7255 0.01144 0.00552 0.0205 0.0258 0.9506
7.000 0.7454 0.01211 0.00628 0.0215 0.0235 0.9535
7.250 0.7714 0.01223 0.00642 0.0215 0.0224 0.9557
7.500 0.7972 0.01258 0.00681 0.0213 0.0209 0.9583
7.750 0.8236 0.01292 0.00718 0.0210 0.0192 0.9610
8.000 0.8456 0.01383 0.00815 0.0211 0.0171 0.9654
8.250 0.8729 0.01424 0.00860 0.0205 0.0163 0.9688
8.500 0.9033 0.01450 0.00889 0.0193 0.0150 0.9712
8.750 0.9336 0.01491 0.00931 0.0180 0.0139 0.9737
9.000 0.9624 0.01554 0.00996 0.0168 0.0129 0.9769
9.250 0.9871 0.01673 0.01125 0.0158 0.0118 0.9814
9.500 1.0151 0.01747 0.01206 0.0146 0.0113 0.9848
9.750 1.0431 0.01810 0.01275 0.0134 0.0106 0.9891
10.000 1.0687 0.01876 0.01344 0.0126 0.0100 0.9976
10.250 1.0899 0.01928 0.01398 0.0128 0.0094 1.0000
10.500 1.0916 0.02000 0.01474 0.0163 0.0091 1.0000
10.750 1.0899 0.02158 0.01642 0.0195 0.0087 1.0000
11.000 1.0953 0.02301 0.01796 0.0216 0.0084 1.0000
11.250 1.1075 0.02390 0.01893 0.0228 0.0082 1.0000
11.500 1.1180 0.02502 0.02014 0.0241 0.0080 1.0000
11.750 1.1285 0.02618 0.02138 0.0254 0.0076 1.0000
12.000 1.1379 0.02746 0.02276 0.0265 0.0075 1.0000
12.250 1.1454 0.02899 0.02443 0.0278 0.0073 1.0000
12.500 1.1546 0.03029 0.02581 0.0288 0.0070 1.0000
12.750 1.1648 0.03146 0.02704 0.0294 0.0068 1.0000
13.000 1.1697 0.03328 0.02898 0.0304 0.0067 1.0000
13.250 1.1791 0.03455 0.03030 0.0309 0.0065 1.0000
13.500 1.1828 0.03648 0.03232 0.0315 0.0063 1.0000
13.750 1.1836 0.03880 0.03477 0.0323 0.0062 1.0000
14.000 1.1818 0.04146 0.03758 0.0329 0.0062 1.0000
14.250 1.1739 0.04488 0.04120 0.0334 0.0062 1.0000
14.500 1.1672 0.04821 0.04467 0.0334 0.0060 1.0000
14.750 1.1383 0.05447 0.05122 0.0330 0.0059 1.0000
15.000 1.1115 0.06100 0.05801 0.0316 0.0057 1.0000
15.250 1.0967 0.06634 0.06353 0.0298 0.0057 1.0000
15.500 1.0826 0.07189 0.06923 0.0275 0.0057 1.0000
15.750 1.0689 0.07787 0.07537 0.0247 0.0057 1.0000
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Polar data table (+)
Polar graphs
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