NACA 6-H-10 AIRFOIL (n6h10-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NACA 6-H-10 AIRFOIL (n6h10-il) Reynolds number: 1,000,000 Max Cl/Cd: 107.27 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n6h10-il-1000000.txt Download as CSV file: xf-n6h10-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.4671 0.09228 0.08964 0.0124 0.6147 0.0065
-7.750 -0.4704 0.08912 0.08649 0.0107 0.6128 0.0066
-7.500 -0.4744 0.08626 0.08363 0.0102 0.6109 0.0066
-7.250 -0.4737 0.08323 0.08058 0.0093 0.6091 0.0068
-7.000 -0.4693 0.07986 0.07717 0.0080 0.6072 0.0069
-6.750 -0.4627 0.07648 0.07375 0.0067 0.6054 0.0071
-6.500 -0.4542 0.07298 0.07022 0.0055 0.6038 0.0072
-6.250 -0.4440 0.06944 0.06663 0.0045 0.6020 0.0074
-6.000 -0.4319 0.06593 0.06306 0.0036 0.6002 0.0076
-5.750 -0.4181 0.06236 0.05942 0.0030 0.5983 0.0079
-5.500 -0.4022 0.05884 0.05581 0.0025 0.5965 0.0082
-5.250 -0.3813 0.05518 0.05201 0.0021 0.5948 0.0086
-5.000 -0.3622 0.05175 0.04844 0.0022 0.5929 0.0086
-4.750 -0.3429 0.04846 0.04503 0.0027 0.5913 0.0087
-4.500 -0.3233 0.04511 0.04156 0.0033 0.5898 0.0087
-4.250 -0.3034 0.04184 0.03815 0.0042 0.5881 0.0087
-4.000 -0.2831 0.03852 0.03467 0.0054 0.5862 0.0088
-3.750 -0.2625 0.03526 0.03124 0.0067 0.5843 0.0088
-3.500 -0.2496 0.03083 0.02657 0.0084 0.5827 0.0093
-3.250 -0.2278 0.02905 0.02466 0.0091 0.5807 0.0097
-3.000 -0.2049 0.02716 0.02260 0.0100 0.5787 0.0105
-2.750 -0.1745 0.02531 0.02050 0.0117 0.5770 0.0125
-2.500 -0.1507 0.02312 0.01805 0.0131 0.5755 0.0126
-2.250 -0.1318 0.01880 0.01334 0.0152 0.5741 0.0134
-2.000 -0.1066 0.01797 0.01246 0.0153 0.5723 0.0141
-1.750 -0.0804 0.01695 0.01130 0.0157 0.5705 0.0151
-1.500 -0.0530 0.01588 0.01005 0.0163 0.5686 0.0164
-1.250 -0.0241 0.01615 0.01019 0.0165 0.5667 0.0180
-1.000 0.0054 0.01240 0.00595 0.0177 0.5651 0.0128
-0.750 0.0340 0.01095 0.00438 0.0178 0.5635 0.0131
-0.500 0.0608 0.01021 0.00363 0.0181 0.5623 0.0141
-0.250 0.0874 0.00979 0.00320 0.0184 0.5609 0.0147
0.000 0.1143 0.00954 0.00296 0.0186 0.5594 0.0160
0.250 0.1405 0.00922 0.00261 0.0190 0.5579 0.0169
0.500 0.1670 0.00898 0.00234 0.0193 0.5564 0.0178
0.750 0.1925 0.00865 0.00196 0.0198 0.5549 0.0201
1.000 0.2192 0.00851 0.00180 0.0201 0.5533 0.0232
1.250 0.2460 0.00840 0.00167 0.0203 0.5517 0.0329
1.500 0.2593 0.00734 0.00169 0.0228 0.5500 0.4512
1.750 0.2587 0.00600 0.00164 0.0289 0.5491 0.8764
2.000 0.2834 0.00608 0.00178 0.0298 0.5479 0.9016
2.250 0.3090 0.00617 0.00191 0.0305 0.5464 0.9155
2.500 0.3344 0.00630 0.00206 0.0313 0.5449 0.9265
2.750 0.3620 0.00637 0.00215 0.0315 0.5431 0.9331
3.000 0.3903 0.00640 0.00217 0.0313 0.5410 0.9351
3.250 0.4180 0.00645 0.00220 0.0313 0.5383 0.9376
3.500 0.4453 0.00650 0.00223 0.0313 0.5349 0.9399
3.750 0.4729 0.00646 0.00221 0.0312 0.5298 0.9413
4.000 0.5013 0.00643 0.00216 0.0310 0.5240 0.9431
4.250 0.5297 0.00645 0.00220 0.0307 0.5204 0.9445
4.500 0.5584 0.00643 0.00221 0.0303 0.5155 0.9450
4.750 0.5868 0.00645 0.00221 0.0300 0.5093 0.9454
5.000 0.6154 0.00645 0.00225 0.0296 0.5035 0.9460
5.250 0.6436 0.00649 0.00227 0.0292 0.4950 0.9465
5.500 0.6721 0.00651 0.00233 0.0288 0.4849 0.9472
5.750 0.7001 0.00658 0.00240 0.0285 0.4706 0.9479
6.000 0.7273 0.00678 0.00251 0.0282 0.4391 0.9488
6.250 0.7471 0.00799 0.00321 0.0283 0.3254 0.9507
6.500 0.7632 0.00945 0.00413 0.0288 0.2037 0.9531
6.750 0.7740 0.01109 0.00521 0.0298 0.0846 0.9562
7.000 0.7977 0.01163 0.00570 0.0296 0.0668 0.9582
7.250 0.8211 0.01217 0.00622 0.0295 0.0569 0.9602
7.500 0.8461 0.01252 0.00662 0.0292 0.0543 0.9621
7.750 0.8697 0.01294 0.00708 0.0291 0.0514 0.9646
8.000 0.8908 0.01349 0.00764 0.0292 0.0478 0.9680
8.250 0.9128 0.01433 0.00856 0.0285 0.0441 0.9714
8.500 0.9424 0.01469 0.00895 0.0270 0.0428 0.9729
8.750 0.9717 0.01540 0.00972 0.0248 0.0411 0.9748
9.000 1.0000 0.01630 0.01066 0.0223 0.0390 0.9772
9.250 1.0245 0.01735 0.01172 0.0203 0.0362 0.9807
9.500 1.0450 0.01895 0.01341 0.0181 0.0338 0.9845
9.750 1.0734 0.01953 0.01404 0.0163 0.0328 0.9859
10.000 1.1012 0.02007 0.01462 0.0147 0.0313 0.9878
10.250 1.1273 0.02074 0.01527 0.0132 0.0278 0.9903
10.500 1.1499 0.02170 0.01623 0.0118 0.0236 0.9936
10.750 1.1714 0.02297 0.01742 0.0100 0.0155 0.9972
11.000 1.1897 0.02480 0.01925 0.0079 0.0110 0.9996
11.250 1.1874 0.02611 0.02057 0.0106 0.0099 1.0000
11.500 1.1819 0.02785 0.02236 0.0134 0.0085 1.0000
11.750 1.1858 0.02919 0.02375 0.0150 0.0084 1.0000
12.000 1.1913 0.03052 0.02515 0.0164 0.0080 1.0000
12.250 1.1957 0.03200 0.02670 0.0177 0.0074 1.0000
12.500 1.2010 0.03348 0.02824 0.0188 0.0069 1.0000
12.750 1.2059 0.03503 0.02985 0.0198 0.0067 1.0000
13.000 1.2092 0.03677 0.03164 0.0208 0.0062 1.0000
13.250 1.2029 0.03944 0.03443 0.0222 0.0056 1.0000
13.500 1.2096 0.04107 0.03612 0.0228 0.0055 1.0000
13.750 1.2120 0.04313 0.03828 0.0235 0.0054 1.0000
14.000 1.2219 0.04456 0.03979 0.0239 0.0050 1.0000
14.250 1.2259 0.04656 0.04188 0.0243 0.0052 1.0000
14.500 1.2297 0.04866 0.04407 0.0247 0.0048 1.0000
14.750 1.2328 0.05085 0.04635 0.0251 0.0047 1.0000
15.000 1.2374 0.05296 0.04854 0.0253 0.0044 1.0000
15.250 1.2392 0.05539 0.05108 0.0255 0.0044 1.0000
15.500 1.2433 0.05765 0.05340 0.0254 0.0041 1.0000
15.750 1.2387 0.06090 0.05680 0.0256 0.0043 1.0000
16.000 1.2425 0.06333 0.05927 0.0252 0.0040 1.0000
16.250 1.2406 0.06647 0.06251 0.0249 0.0038 1.0000
16.500 1.2326 0.07043 0.06659 0.0245 0.0037 1.0000
16.750 1.2275 0.07415 0.07043 0.0239 0.0037 1.0000
17.000 1.2178 0.07863 0.07506 0.0232 0.0037 1.0000
17.250 1.2027 0.08395 0.08054 0.0220 0.0036 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 6-H-10 AIRFOIL (n6h10-il)