NACA 8-H-12 AIRFOIL (n8h12-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 8-H-12 AIRFOIL (n8h12-il) Reynolds number: 500,000 Max Cl/Cd: 98.64 at α=6.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n8h12-il-500000-n5.txt Download as CSV file: xf-n8h12-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 8-H-12 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.5368 0.09060 0.08753 0.0082 0.6654 0.0105
-9.500 -0.5487 0.08430 0.08124 0.0046 0.6611 0.0107
-9.250 -0.8087 0.02816 0.02327 -0.0083 0.6757 0.0115
-9.000 -0.7967 0.02494 0.01949 -0.0067 0.6697 0.0119
-8.750 -0.7765 0.02365 0.01800 -0.0058 0.6629 0.0121
-8.500 -0.7534 0.02295 0.01719 -0.0052 0.6568 0.0124
-8.250 -0.7293 0.02239 0.01654 -0.0047 0.6502 0.0126
-8.000 -0.7053 0.02177 0.01577 -0.0041 0.6438 0.0130
-7.750 -0.6814 0.02096 0.01480 -0.0035 0.6372 0.0135
-7.500 -0.6576 0.02004 0.01365 -0.0029 0.6308 0.0141
-7.250 -0.6338 0.01898 0.01233 -0.0022 0.6252 0.0146
-7.000 -0.6094 0.01808 0.01123 -0.0015 0.6190 0.0150
-6.750 -0.5844 0.01746 0.01052 -0.0011 0.6129 0.0155
-6.500 -0.5585 0.01699 0.00998 -0.0007 0.6066 0.0159
-6.250 -0.5326 0.01652 0.00940 -0.0003 0.6000 0.0164
-6.000 -0.5065 0.01602 0.00879 0.0001 0.5942 0.0171
-5.750 -0.4801 0.01554 0.00819 0.0005 0.5883 0.0179
-5.500 -0.4536 0.01512 0.00762 0.0009 0.5825 0.0185
-5.250 -0.4280 0.01450 0.00696 0.0013 0.5771 0.0192
-5.000 -0.4015 0.01411 0.00652 0.0016 0.5708 0.0198
-4.500 -0.3482 0.01338 0.00566 0.0023 0.5592 0.0214
-4.250 -0.3212 0.01310 0.00529 0.0026 0.5535 0.0223
-4.000 -0.2950 0.01268 0.00480 0.0030 0.5485 0.0232
-3.750 -0.2684 0.01232 0.00443 0.0033 0.5434 0.0240
-3.500 -0.2414 0.01205 0.00412 0.0036 0.5381 0.0249
-3.000 -0.1870 0.01161 0.00357 0.0041 0.5280 0.0272
-2.750 -0.1600 0.01135 0.00327 0.0043 0.5226 0.0283
-2.500 -0.1330 0.01114 0.00303 0.0046 0.5177 0.0298
-2.250 -0.1055 0.01096 0.00283 0.0048 0.5133 0.0316
-2.000 -0.0779 0.01081 0.00265 0.0049 0.5087 0.0337
-1.750 -0.0506 0.01064 0.00247 0.0052 0.5043 0.0378
-1.500 -0.0231 0.01050 0.00233 0.0053 0.5003 0.0449
-1.250 0.0044 0.01034 0.00222 0.0055 0.4962 0.0599
-1.000 0.0318 0.01019 0.00212 0.0056 0.4919 0.0786
-0.750 0.0590 0.01005 0.00203 0.0058 0.4877 0.1026
-0.500 0.0857 0.00981 0.00195 0.0060 0.4839 0.1543
-0.250 0.1117 0.00949 0.00189 0.0062 0.4800 0.2346
0.000 0.1345 0.00885 0.00181 0.0070 0.4765 0.4106
0.250 0.1487 0.00779 0.00174 0.0096 0.4733 0.6960
0.500 0.2176 0.00732 0.00199 0.0015 0.4693 0.9380
0.750 0.2553 0.00750 0.00213 -0.0002 0.4659 0.9633
1.000 0.3034 0.00770 0.00227 -0.0043 0.4622 0.9778
1.250 0.3519 0.00786 0.00236 -0.0085 0.4584 0.9871
1.500 0.3922 0.00794 0.00237 -0.0112 0.4549 0.9912
1.750 0.4266 0.00799 0.00239 -0.0126 0.4519 0.9941
2.000 0.4608 0.00801 0.00239 -0.0140 0.4490 0.9959
2.250 0.4945 0.00803 0.00239 -0.0153 0.4460 0.9978
2.500 0.5278 0.00808 0.00241 -0.0165 0.4432 0.9994
2.750 0.5567 0.00814 0.00244 -0.0168 0.4406 1.0000
3.000 0.5828 0.00820 0.00249 -0.0166 0.4380 1.0000
3.250 0.6090 0.00825 0.00256 -0.0163 0.4354 1.0000
3.500 0.6352 0.00831 0.00262 -0.0160 0.4328 1.0000
3.750 0.6614 0.00838 0.00270 -0.0157 0.4303 1.0000
4.000 0.6876 0.00847 0.00279 -0.0155 0.4277 1.0000
4.250 0.7136 0.00857 0.00289 -0.0152 0.4252 1.0000
4.500 0.7398 0.00868 0.00300 -0.0150 0.4228 1.0000
4.750 0.7660 0.00876 0.00312 -0.0147 0.4203 1.0000
5.000 0.7923 0.00884 0.00325 -0.0145 0.4164 1.0000
5.250 0.8183 0.00895 0.00335 -0.0142 0.4111 1.0000
5.500 0.8444 0.00906 0.00348 -0.0140 0.4059 1.0000
5.750 0.8705 0.00915 0.00361 -0.0138 0.3977 1.0000
6.000 0.8965 0.00927 0.00373 -0.0136 0.3868 1.0000
6.250 0.9222 0.00943 0.00387 -0.0134 0.3760 1.0000
6.500 0.9479 0.00961 0.00404 -0.0132 0.3578 1.0000
6.750 0.9728 0.00993 0.00426 -0.0130 0.3290 1.0000
7.000 0.9963 0.01049 0.00463 -0.0129 0.2858 1.0000
7.250 1.0161 0.01163 0.00544 -0.0127 0.2205 1.0000
7.500 1.0353 0.01273 0.00628 -0.0123 0.1705 1.0000
7.750 1.0538 0.01377 0.00711 -0.0118 0.1294 1.0000
8.000 1.0720 0.01470 0.00789 -0.0112 0.0994 1.0000
8.250 1.0893 0.01560 0.00868 -0.0104 0.0761 1.0000
8.750 1.1217 0.01731 0.01027 -0.0086 0.0483 1.0000
9.250 1.1496 0.01903 0.01199 -0.0065 0.0344 1.0000
9.500 1.1598 0.02011 0.01310 -0.0054 0.0301 1.0000
9.750 1.1602 0.02136 0.01437 -0.0030 0.0275 1.0000
10.000 1.1629 0.02266 0.01570 -0.0010 0.0254 1.0000
10.250 1.1686 0.02401 0.01710 0.0002 0.0240 1.0000
10.500 1.1740 0.02553 0.01865 0.0013 0.0227 1.0000
10.750 1.1783 0.02724 0.02041 0.0021 0.0214 1.0000
11.000 1.1841 0.02891 0.02214 0.0028 0.0209 1.0000
11.250 1.1903 0.03060 0.02389 0.0034 0.0201 1.0000
11.500 1.1955 0.03243 0.02579 0.0039 0.0195 1.0000
11.750 1.2000 0.03435 0.02777 0.0043 0.0189 1.0000
12.000 1.2040 0.03636 0.02984 0.0047 0.0184 1.0000
12.250 1.2074 0.03845 0.03200 0.0050 0.0182 1.0000
12.500 1.2099 0.04068 0.03429 0.0053 0.0178 1.0000
12.750 1.2111 0.04306 0.03672 0.0055 0.0175 1.0000
13.000 1.2100 0.04574 0.03946 0.0056 0.0170 1.0000
13.250 1.2115 0.04823 0.04202 0.0056 0.0169 1.0000
13.500 1.2144 0.05061 0.04448 0.0055 0.0167 1.0000
13.750 1.2175 0.05304 0.04698 0.0054 0.0165 1.0000
14.000 1.2205 0.05552 0.04955 0.0052 0.0162 1.0000
14.250 1.2224 0.05813 0.05223 0.0050 0.0160 1.0000
14.500 1.2250 0.06073 0.05490 0.0047 0.0157 1.0000
14.750 1.2278 0.06331 0.05756 0.0044 0.0154 1.0000
15.000 1.2293 0.06609 0.06040 0.0041 0.0153 1.0000
15.250 1.2309 0.06889 0.06327 0.0036 0.0150 1.0000
15.500 1.2324 0.07172 0.06618 0.0032 0.0148 1.0000
15.750 1.2334 0.07464 0.06916 0.0027 0.0147 1.0000
16.000 1.2349 0.07756 0.07216 0.0021 0.0145 1.0000
16.250 1.2352 0.08063 0.07530 0.0015 0.0143 1.0000
16.500 1.2358 0.08370 0.07844 0.0008 0.0142 1.0000
16.750 1.2356 0.08691 0.08170 0.0000 0.0139 1.0000
17.000 1.2351 0.09015 0.08501 -0.0007 0.0138 1.0000
17.250 1.2336 0.09355 0.08847 -0.0015 0.0136 1.0000
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