ONERA NACA CAMBRE AIRFOIL (ncambre-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: ONERA NACA CAMBRE AIRFOIL (ncambre-il) Reynolds number: 500,000 Max Cl/Cd: 72.96 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ncambre-il-500000-n5.txt Download as CSV file: xf-ncambre-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: ONERA NACA CAMBRE 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
------ -------- --------- --------- -------- -------- --------
-11.000 -0.6554 0.08719 0.08474 -0.0135 1.0000 0.0143
-10.750 -0.7033 0.06867 0.06613 -0.0301 1.0000 0.0141
-10.500 -0.7396 0.06052 0.05779 -0.0348 1.0000 0.0141
-10.250 -0.7720 0.05467 0.05174 -0.0338 1.0000 0.0141
-10.000 -0.8073 0.04658 0.04324 -0.0310 1.0000 0.0142
-9.750 -0.8699 0.03123 0.02650 -0.0239 1.0000 0.0149
-9.500 -0.8611 0.02884 0.02385 -0.0218 1.0000 0.0150
-9.250 -0.8469 0.02731 0.02217 -0.0202 1.0000 0.0152
-9.000 -0.8311 0.02609 0.02082 -0.0186 1.0000 0.0153
-8.750 -0.8147 0.02499 0.01960 -0.0169 1.0000 0.0154
-8.500 -0.7980 0.02394 0.01842 -0.0153 1.0000 0.0156
-8.250 -0.7809 0.02292 0.01727 -0.0136 1.0000 0.0158
-8.000 -0.7636 0.02192 0.01614 -0.0119 1.0000 0.0160
-7.750 -0.7460 0.02094 0.01504 -0.0102 1.0000 0.0162
-7.500 -0.7278 0.02000 0.01398 -0.0085 0.9999 0.0165
-7.250 -0.6967 0.01892 0.01275 -0.0096 0.9974 0.0168
-7.000 -0.6643 0.01792 0.01162 -0.0108 0.9947 0.0171
-6.750 -0.6330 0.01702 0.01061 -0.0118 0.9915 0.0175
-6.500 -0.6016 0.01623 0.00972 -0.0128 0.9879 0.0179
-6.250 -0.5680 0.01560 0.00900 -0.0141 0.9848 0.0182
-6.000 -0.5381 0.01481 0.00819 -0.0149 0.9796 0.0188
-5.750 -0.5076 0.01429 0.00766 -0.0156 0.9743 0.0192
-5.500 -0.4767 0.01379 0.00714 -0.0164 0.9698 0.0197
-5.250 -0.4486 0.01332 0.00665 -0.0165 0.9621 0.0201
-5.000 -0.4186 0.01285 0.00615 -0.0170 0.9549 0.0207
-4.750 -0.3910 0.01242 0.00568 -0.0169 0.9447 0.0213
-4.500 -0.3636 0.01203 0.00525 -0.0168 0.9335 0.0219
-4.250 -0.3368 0.01164 0.00482 -0.0165 0.9205 0.0225
-4.000 -0.3107 0.01126 0.00441 -0.0160 0.9047 0.0233
-3.750 -0.2847 0.01101 0.00412 -0.0155 0.8874 0.0242
-3.500 -0.2587 0.01080 0.00385 -0.0150 0.8693 0.0254
-3.250 -0.2325 0.01060 0.00359 -0.0145 0.8512 0.0266
-3.000 -0.2069 0.01037 0.00329 -0.0138 0.8314 0.0279
-2.750 -0.1809 0.01021 0.00306 -0.0133 0.8096 0.0294
-2.500 -0.1547 0.01010 0.00284 -0.0128 0.7857 0.0312
-2.250 -0.1284 0.00998 0.00264 -0.0124 0.7617 0.0337
-2.000 -0.1021 0.00989 0.00247 -0.0119 0.7359 0.0372
-1.750 -0.0757 0.00981 0.00230 -0.0115 0.7111 0.0419
-1.500 -0.0494 0.00974 0.00215 -0.0111 0.6845 0.0508
-1.250 -0.0234 0.00960 0.00202 -0.0107 0.6569 0.0821
-0.750 0.0201 0.00798 0.00162 -0.0092 0.6117 0.4744
-0.500 0.0446 0.00771 0.00157 -0.0085 0.5906 0.5585
-0.250 0.0693 0.00753 0.00155 -0.0079 0.5696 0.6285
0.000 0.0944 0.00739 0.00155 -0.0072 0.5486 0.6864
0.250 0.1192 0.00728 0.00159 -0.0064 0.5287 0.7431
0.500 0.1447 0.00725 0.00162 -0.0057 0.5099 0.7859
0.750 0.1703 0.00728 0.00168 -0.0051 0.4891 0.8205
1.000 0.1958 0.00735 0.00175 -0.0043 0.4666 0.8521
1.250 0.2221 0.00745 0.00181 -0.0038 0.4450 0.8747
1.500 0.2490 0.00755 0.00188 -0.0035 0.4268 0.8915
2.000 0.3039 0.00785 0.00203 -0.0030 0.3827 0.9184
2.250 0.3319 0.00801 0.00212 -0.0030 0.3634 0.9293
2.500 0.3596 0.00820 0.00221 -0.0029 0.3446 0.9396
2.750 0.3892 0.00840 0.00232 -0.0033 0.3234 0.9474
3.000 0.4170 0.00860 0.00242 -0.0033 0.2999 0.9561
3.250 0.4484 0.00882 0.00254 -0.0041 0.2812 0.9611
3.500 0.4785 0.00900 0.00267 -0.0046 0.2673 0.9676
3.750 0.5101 0.00920 0.00280 -0.0055 0.2528 0.9733
4.000 0.5431 0.00942 0.00296 -0.0067 0.2373 0.9783
4.250 0.5740 0.00964 0.00311 -0.0075 0.2240 0.9836
4.500 0.6089 0.00986 0.00327 -0.0092 0.2119 0.9860
5.000 0.6763 0.01031 0.00363 -0.0120 0.1887 0.9927
5.250 0.7097 0.01053 0.00382 -0.0134 0.1784 0.9952
5.500 0.7431 0.01078 0.00403 -0.0148 0.1680 0.9976
5.750 0.7763 0.01101 0.00423 -0.0162 0.1580 0.9999
6.000 0.7992 0.01123 0.00443 -0.0153 0.1502 1.0000
6.250 0.8215 0.01147 0.00465 -0.0144 0.1426 1.0000
6.500 0.8441 0.01170 0.00487 -0.0134 0.1365 1.0000
6.750 0.8665 0.01195 0.00511 -0.0125 0.1303 1.0000
7.000 0.8888 0.01223 0.00537 -0.0115 0.1241 1.0000
7.250 0.9113 0.01249 0.00564 -0.0106 0.1172 1.0000
7.500 0.9334 0.01280 0.00593 -0.0096 0.1103 1.0000
7.750 0.9558 0.01310 0.00621 -0.0087 0.1031 1.0000
8.000 0.9780 0.01343 0.00652 -0.0078 0.0966 1.0000
8.250 1.0006 0.01375 0.00684 -0.0070 0.0914 1.0000
8.500 1.0230 0.01412 0.00720 -0.0062 0.0869 1.0000
8.750 1.0464 0.01440 0.00753 -0.0055 0.0837 1.0000
9.000 1.0693 0.01475 0.00789 -0.0048 0.0797 1.0000
9.250 1.0916 0.01516 0.00830 -0.0041 0.0754 1.0000
9.500 1.1149 0.01548 0.00867 -0.0035 0.0718 1.0000
9.750 1.1371 0.01591 0.00910 -0.0027 0.0662 1.0000
10.000 1.1594 0.01633 0.00954 -0.0020 0.0608 1.0000
10.250 1.1807 0.01683 0.01003 -0.0013 0.0538 1.0000
10.500 1.2010 0.01743 0.01059 -0.0004 0.0441 1.0000
10.750 1.2193 0.01819 0.01130 0.0008 0.0351 1.0000
11.000 1.2376 0.01892 0.01204 0.0019 0.0300 1.0000
11.250 1.2554 0.01967 0.01282 0.0031 0.0267 1.0000
11.500 1.2735 0.02035 0.01355 0.0042 0.0244 1.0000
11.750 1.2898 0.02114 0.01439 0.0055 0.0226 1.0000
12.000 1.3064 0.02186 0.01519 0.0068 0.0215 1.0000
12.250 1.3215 0.02262 0.01602 0.0083 0.0204 1.0000
12.500 1.3332 0.02345 0.01691 0.0102 0.0194 1.0000
12.750 1.3427 0.02443 0.01796 0.0122 0.0186 1.0000
13.000 1.3522 0.02545 0.01907 0.0141 0.0179 1.0000
13.250 1.3620 0.02649 0.02020 0.0157 0.0174 1.0000
13.500 1.3703 0.02770 0.02151 0.0172 0.0170 1.0000
13.750 1.3773 0.02908 0.02298 0.0185 0.0165 1.0000
14.000 1.3828 0.03067 0.02468 0.0196 0.0161 1.0000
14.250 1.3865 0.03255 0.02666 0.0204 0.0157 1.0000
14.500 1.3880 0.03480 0.02902 0.0209 0.0154 1.0000
14.750 1.3866 0.03756 0.03189 0.0208 0.0150 1.0000
15.000 1.3817 0.04096 0.03541 0.0202 0.0147 1.0000
15.250 1.3804 0.04417 0.03876 0.0193 0.0145 1.0000
15.500 1.3767 0.04789 0.04261 0.0179 0.0143 1.0000
15.750 1.3697 0.05222 0.04708 0.0161 0.0141 1.0000
16.000 1.3597 0.05716 0.05217 0.0139 0.0140 1.0000
16.250 1.3461 0.06273 0.05789 0.0113 0.0139 1.0000
16.500 1.3295 0.06888 0.06419 0.0084 0.0138 1.0000
16.750 1.3101 0.07554 0.07100 0.0053 0.0137 1.0000
17.000 1.2888 0.08270 0.07830 0.0020 0.0137 1.0000
17.250 1.2663 0.09015 0.08588 -0.0015 0.0137 1.0000
|
Polar data table (+)
Polar graphs
<< Back to ONERA NACA CAMBRE AIRFOIL (ncambre-il)