NACA 11-H-09 AIRFOIL (n11h9-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NACA 11-H-09 AIRFOIL (n11h9-il) Reynolds number: 1,000,000 Max Cl/Cd: 138.89 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n11h9-il-1000000.txt Download as CSV file: xf-n11h9-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 11-H-09 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 -0.3751 0.09035 0.08778 -0.0040 0.6628 0.0066
-7.750 -0.3790 0.08736 0.08481 -0.0056 0.6587 0.0066
-7.500 -0.3827 0.08447 0.08191 -0.0065 0.6548 0.0066
-7.250 -0.3336 0.06956 0.06715 -0.0031 0.6400 0.0068
-7.000 -0.3358 0.06651 0.06410 -0.0032 0.6366 0.0069
-6.750 -0.3357 0.06296 0.06054 -0.0044 0.6332 0.0069
-6.500 -0.3311 0.05949 0.05705 -0.0057 0.6297 0.0072
-6.250 -0.3249 0.05580 0.05332 -0.0072 0.6263 0.0073
-6.000 -0.3166 0.05198 0.04945 -0.0087 0.6232 0.0075
-5.750 -0.3062 0.04814 0.04558 -0.0100 0.6200 0.0077
-5.500 -0.2938 0.04429 0.04168 -0.0112 0.6166 0.0082
-5.250 -0.2795 0.04050 0.03780 -0.0123 0.6133 0.0086
-5.000 -0.2596 0.03679 0.03396 -0.0134 0.6102 0.0092
-4.750 -0.2394 0.03319 0.03024 -0.0141 0.6073 0.0093
-4.500 -0.2205 0.02980 0.02675 -0.0143 0.6040 0.0093
-4.250 -0.2014 0.02657 0.02339 -0.0141 0.6007 0.0093
-4.000 -0.1816 0.02347 0.02016 -0.0138 0.5975 0.0093
-3.750 -0.1615 0.02059 0.01713 -0.0132 0.5944 0.0094
-3.500 -0.1460 0.01646 0.01284 -0.0124 0.5918 0.0096
-3.250 -0.1265 0.01439 0.01067 -0.0120 0.5887 0.0098
-3.000 -0.1055 0.01264 0.00881 -0.0114 0.5856 0.0101
-2.750 -0.0834 0.01097 0.00698 -0.0106 0.5827 0.0104
-2.500 -0.0605 0.00943 0.00526 -0.0097 0.5798 0.0110
-2.250 -0.0329 0.00813 0.00374 -0.0081 0.5774 0.0126
-2.000 -0.0078 0.00684 0.00219 -0.0066 0.5746 0.0128
-1.250 -0.0109 0.12948 0.12646 0.0061 0.5739 0.0193
-0.750 0.1133 0.01172 0.00599 -0.0019 0.5633 0.0190
-0.250 0.1766 0.01008 0.00373 -0.0010 0.5595 0.0143
0.000 0.2034 0.00970 0.00331 -0.0007 0.5568 0.0156
0.250 0.2294 0.00927 0.00282 -0.0002 0.5540 0.0163
0.500 0.2564 0.00918 0.00269 0.0001 0.5507 0.0173
0.750 0.2804 0.00853 0.00202 0.0010 0.5475 0.0188
1.000 0.3070 0.00835 0.00184 0.0014 0.5437 0.0210
1.250 0.3337 0.00824 0.00170 0.0017 0.5404 0.0230
1.500 0.3598 0.00811 0.00152 0.0021 0.5372 0.0278
1.750 0.3867 0.00799 0.00142 0.0024 0.5348 0.0370
2.000 0.4135 0.00786 0.00136 0.0026 0.5320 0.0670
2.250 0.5232 0.00621 0.00202 -0.0162 0.5280 0.9874
2.500 0.5912 0.00634 0.00207 -0.0251 0.5238 0.9959
2.750 0.6390 0.00626 0.00196 -0.0297 0.5198 0.9995
3.000 0.6683 0.00625 0.00196 -0.0301 0.5159 1.0000
3.250 0.6941 0.00628 0.00197 -0.0297 0.5118 1.0000
3.500 0.7196 0.00634 0.00200 -0.0293 0.5074 1.0000
3.750 0.7457 0.00635 0.00203 -0.0291 0.5033 1.0000
4.000 0.7716 0.00638 0.00207 -0.0287 0.4990 1.0000
4.250 0.7972 0.00644 0.00213 -0.0284 0.4946 1.0000
4.500 0.8231 0.00648 0.00219 -0.0281 0.4904 1.0000
4.750 0.8491 0.00651 0.00224 -0.0279 0.4846 1.0000
5.000 0.8748 0.00658 0.00230 -0.0275 0.4769 1.0000
5.250 0.9004 0.00665 0.00237 -0.0272 0.4644 1.0000
5.500 0.9260 0.00673 0.00244 -0.0270 0.4489 1.0000
5.750 0.9514 0.00685 0.00253 -0.0267 0.4310 1.0000
6.000 0.9762 0.00708 0.00268 -0.0263 0.4028 1.0000
6.250 0.9994 0.00754 0.00295 -0.0259 0.3574 1.0000
6.500 1.0219 0.00814 0.00334 -0.0255 0.3106 1.0000
6.750 1.0448 0.00869 0.00372 -0.0251 0.2713 1.0000
7.000 1.0672 0.00932 0.00416 -0.0248 0.2291 1.0000
7.250 1.0887 0.01008 0.00471 -0.0244 0.1844 1.0000
7.500 1.1096 0.01090 0.00531 -0.0240 0.1416 1.0000
7.750 1.1290 0.01190 0.00605 -0.0235 0.0947 1.0000
8.000 1.1471 0.01301 0.00691 -0.0229 0.0532 1.0000
8.250 1.1649 0.01407 0.00780 -0.0223 0.0272 1.0000
8.500 1.1848 0.01475 0.00850 -0.0216 0.0220 1.0000
8.750 1.2053 0.01529 0.00909 -0.0210 0.0207 1.0000
9.000 1.2244 0.01593 0.00978 -0.0203 0.0194 1.0000
9.250 1.2419 0.01670 0.01060 -0.0194 0.0182 1.0000
9.500 1.2575 0.01762 0.01160 -0.0186 0.0173 1.0000
9.750 1.2675 0.01895 0.01305 -0.0174 0.0164 1.0000
10.000 1.2759 0.02015 0.01433 -0.0159 0.0160 1.0000
10.250 1.2779 0.02110 0.01534 -0.0131 0.0159 1.0000
10.500 1.2794 0.02215 0.01645 -0.0104 0.0156 1.0000
10.750 1.2837 0.02334 0.01771 -0.0084 0.0153 1.0000
11.000 1.2884 0.02465 0.01909 -0.0067 0.0149 1.0000
11.250 1.2944 0.02597 0.02047 -0.0053 0.0144 1.0000
11.500 1.2983 0.02753 0.02210 -0.0039 0.0140 1.0000
11.750 1.3063 0.02881 0.02342 -0.0029 0.0134 1.0000
12.000 1.3103 0.03048 0.02513 -0.0018 0.0128 1.0000
12.250 1.3126 0.03233 0.02705 -0.0008 0.0125 1.0000
12.500 1.3098 0.03470 0.02950 0.0004 0.0119 1.0000
12.750 1.3056 0.03727 0.03217 0.0015 0.0117 1.0000
13.000 1.3061 0.03943 0.03442 0.0024 0.0112 1.0000
13.250 1.3079 0.04156 0.03665 0.0031 0.0111 1.0000
13.500 1.3194 0.04283 0.03797 0.0033 0.0107 1.0000
13.750 1.3244 0.04473 0.03994 0.0037 0.0104 1.0000
14.000 1.3325 0.04640 0.04167 0.0039 0.0099 1.0000
14.250 1.3340 0.04877 0.04412 0.0042 0.0096 1.0000
14.500 1.3394 0.05080 0.04621 0.0043 0.0092 1.0000
14.750 1.3448 0.05289 0.04833 0.0042 0.0088 1.0000
15.000 1.3440 0.05570 0.05120 0.0042 0.0085 1.0000
15.250 1.3299 0.06005 0.05566 0.0042 0.0078 1.0000
15.500 1.3334 0.06251 0.05821 0.0040 0.0077 1.0000
15.750 1.3319 0.06561 0.06140 0.0037 0.0076 1.0000
16.000 1.3489 0.06657 0.06242 0.0032 0.0071 1.0000
16.250 1.3519 0.06923 0.06515 0.0027 0.0066 1.0000
16.500 1.3526 0.07222 0.06820 0.0022 0.0063 1.0000
16.750 1.3529 0.07528 0.07135 0.0015 0.0062 1.0000
17.000 1.3514 0.07869 0.07482 0.0007 0.0059 1.0000
17.250 1.3468 0.08260 0.07881 -0.0003 0.0057 1.0000
17.500 1.3396 0.08692 0.08321 -0.0015 0.0055 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 11-H-09 AIRFOIL (n11h9-il)