ONERA OA212 AIRFOIL (oa212-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: ONERA OA212 AIRFOIL (oa212-il) Reynolds number: 500,000 Max Cl/Cd: 76.34 at α=7.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-oa212-il-500000-n5.txt Download as CSV file: xf-oa212-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: ONERA OA212 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
------ -------- --------- --------- -------- -------- --------
-14.250 -1.1587 0.05953 0.05642 -0.0113 0.9174 0.0177
-14.000 -1.1977 0.04995 0.04649 -0.0175 0.9138 0.0176
-13.750 -1.2170 0.04524 0.04152 -0.0177 0.9098 0.0177
-13.500 -1.2296 0.04193 0.03797 -0.0160 0.9057 0.0178
-13.250 -1.2397 0.03950 0.03532 -0.0129 0.9020 0.0178
-13.000 -1.2391 0.03733 0.03291 -0.0107 0.8988 0.0179
-12.750 -1.2336 0.03527 0.03069 -0.0090 0.8952 0.0181
-12.500 -1.2237 0.03358 0.02886 -0.0074 0.8919 0.0182
-12.250 -1.2106 0.03218 0.02733 -0.0061 0.8889 0.0184
-12.000 -1.1957 0.03094 0.02596 -0.0047 0.8862 0.0186
-11.750 -1.1782 0.02978 0.02469 -0.0038 0.8833 0.0188
-11.500 -1.1588 0.02868 0.02347 -0.0031 0.8799 0.0190
-11.250 -1.1390 0.02763 0.02229 -0.0023 0.8766 0.0192
-11.000 -1.1189 0.02662 0.02115 -0.0015 0.8733 0.0195
-10.750 -1.0984 0.02565 0.02004 -0.0006 0.8705 0.0198
-10.500 -1.0770 0.02471 0.01896 0.0002 0.8678 0.0201
-10.250 -1.0536 0.02378 0.01790 0.0006 0.8642 0.0204
-10.000 -1.0302 0.02292 0.01689 0.0010 0.8605 0.0207
-9.750 -1.0065 0.02212 0.01595 0.0015 0.8571 0.0210
-9.500 -0.9831 0.02133 0.01504 0.0021 0.8541 0.0213
-9.250 -0.9594 0.02061 0.01426 0.0027 0.8512 0.0217
-9.000 -0.9332 0.02001 0.01362 0.0028 0.8472 0.0222
-8.750 -0.9071 0.01946 0.01300 0.0029 0.8433 0.0226
-8.500 -0.8813 0.01890 0.01237 0.0032 0.8397 0.0231
-8.250 -0.8558 0.01834 0.01172 0.0036 0.8363 0.0236
-8.000 -0.8297 0.01780 0.01108 0.0039 0.8329 0.0241
-7.750 -0.8023 0.01729 0.01048 0.0039 0.8289 0.0246
-7.500 -0.7756 0.01671 0.00986 0.0040 0.8247 0.0252
-7.250 -0.7489 0.01624 0.00935 0.0042 0.8208 0.0258
-7.000 -0.7222 0.01583 0.00888 0.0044 0.8173 0.0264
-6.750 -0.6942 0.01543 0.00844 0.0044 0.8134 0.0272
-6.500 -0.6661 0.01505 0.00801 0.0043 0.8089 0.0281
-6.250 -0.6384 0.01467 0.00757 0.0044 0.8046 0.0289
-6.000 -0.6110 0.01428 0.00716 0.0045 0.8007 0.0300
-5.750 -0.5828 0.01395 0.00680 0.0044 0.7969 0.0310
-5.500 -0.5541 0.01363 0.00645 0.0043 0.7923 0.0322
-5.250 -0.5257 0.01332 0.00609 0.0042 0.7876 0.0333
-5.000 -0.4976 0.01300 0.00576 0.0042 0.7835 0.0348
-4.750 -0.4690 0.01274 0.00548 0.0041 0.7796 0.0365
-4.500 -0.4398 0.01248 0.00520 0.0039 0.7748 0.0384
-4.250 -0.4111 0.01221 0.00493 0.0038 0.7697 0.0405
-4.000 -0.3826 0.01199 0.00466 0.0038 0.7642 0.0428
-3.750 -0.3533 0.01174 0.00443 0.0035 0.7561 0.0454
-3.500 -0.3245 0.01154 0.00419 0.0034 0.7477 0.0488
-3.250 -0.2952 0.01133 0.00398 0.0032 0.7396 0.0524
-3.000 -0.2660 0.01116 0.00378 0.0030 0.7310 0.0564
-2.750 -0.2365 0.01097 0.00360 0.0028 0.7226 0.0611
-2.500 -0.2070 0.01082 0.00343 0.0025 0.7147 0.0661
-2.250 -0.1775 0.01066 0.00329 0.0022 0.7079 0.0718
-2.000 -0.1477 0.01051 0.00315 0.0019 0.7006 0.0776
-1.750 -0.1181 0.01039 0.00302 0.0016 0.6936 0.0843
-1.500 -0.0882 0.01024 0.00291 0.0012 0.6855 0.0917
-1.250 -0.0583 0.01014 0.00278 0.0008 0.6770 0.0997
-1.000 -0.0283 0.01000 0.00269 0.0004 0.6689 0.1098
-0.750 0.0016 0.00989 0.00259 0.0000 0.6594 0.1219
-0.500 0.0317 0.00976 0.00251 -0.0004 0.6490 0.1371
-0.250 0.0618 0.00964 0.00243 -0.0009 0.6384 0.1585
0.000 0.0918 0.00947 0.00236 -0.0014 0.6268 0.1943
0.250 0.1218 0.00920 0.00230 -0.0020 0.6137 0.2650
0.500 0.1516 0.00880 0.00223 -0.0027 0.5985 0.3781
0.750 0.1712 0.00726 0.00236 -0.0011 0.5841 0.8440
1.000 0.1984 0.00740 0.00246 -0.0007 0.5669 0.8789
1.250 0.2270 0.00756 0.00254 -0.0006 0.5447 0.8950
1.500 0.2556 0.00775 0.00260 -0.0006 0.5187 0.9058
1.750 0.2843 0.00797 0.00269 -0.0007 0.4908 0.9170
2.000 0.3100 0.00820 0.00281 0.0000 0.4630 0.9278
2.250 0.3351 0.00842 0.00291 0.0008 0.4367 0.9373
2.500 0.3597 0.00861 0.00300 0.0017 0.4159 0.9476
2.750 0.3827 0.00873 0.00306 0.0030 0.4001 0.9575
3.000 0.4096 0.00885 0.00310 0.0033 0.3855 0.9626
3.250 0.4396 0.00899 0.00317 0.0027 0.3725 0.9640
3.500 0.4696 0.00913 0.00326 0.0021 0.3619 0.9654
3.750 0.4995 0.00926 0.00334 0.0016 0.3516 0.9670
4.000 0.5294 0.00941 0.00344 0.0010 0.3426 0.9683
4.250 0.5594 0.00955 0.00355 0.0004 0.3339 0.9695
4.500 0.5891 0.00971 0.00367 -0.0002 0.3261 0.9708
5.000 0.6486 0.01001 0.00392 -0.0013 0.3110 0.9735
5.250 0.6786 0.01014 0.00405 -0.0019 0.3039 0.9746
5.500 0.7084 0.01032 0.00420 -0.0025 0.2962 0.9757
5.750 0.7383 0.01047 0.00436 -0.0031 0.2891 0.9769
6.000 0.7679 0.01067 0.00453 -0.0037 0.2803 0.9782
6.250 0.7974 0.01086 0.00471 -0.0043 0.2697 0.9795
6.500 0.8265 0.01113 0.00490 -0.0049 0.2525 0.9812
6.750 0.8553 0.01141 0.00511 -0.0055 0.2358 0.9830
7.000 0.8848 0.01169 0.00535 -0.0062 0.2223 0.9842
7.250 0.9144 0.01200 0.00560 -0.0070 0.2077 0.9854
7.500 0.9436 0.01236 0.00589 -0.0078 0.1906 0.9867
7.750 0.9721 0.01280 0.00625 -0.0085 0.1709 0.9882
8.000 0.9999 0.01332 0.00665 -0.0092 0.1484 0.9900
8.250 1.0267 0.01395 0.00714 -0.0098 0.1235 0.9922
8.750 1.0776 0.01537 0.00830 -0.0107 0.0786 1.0000
9.000 1.1003 0.01606 0.00890 -0.0107 0.0630 1.0000
9.250 1.1231 0.01673 0.00952 -0.0106 0.0519 1.0000
9.500 1.1460 0.01737 0.01012 -0.0105 0.0445 1.0000
9.750 1.1686 0.01800 0.01074 -0.0104 0.0394 1.0000
10.000 1.1907 0.01865 0.01139 -0.0103 0.0355 1.0000
10.250 1.2125 0.01929 0.01206 -0.0101 0.0327 1.0000
10.750 1.2524 0.02075 0.01357 -0.0096 0.0284 1.0000
11.000 1.2684 0.02155 0.01441 -0.0088 0.0268 1.0000
11.250 1.2831 0.02246 0.01535 -0.0079 0.0257 1.0000
11.500 1.2990 0.02332 0.01627 -0.0072 0.0247 1.0000
11.750 1.3140 0.02429 0.01729 -0.0065 0.0237 1.0000
12.000 1.3276 0.02537 0.01842 -0.0058 0.0228 1.0000
12.250 1.3391 0.02665 0.01974 -0.0050 0.0220 1.0000
12.500 1.3521 0.02783 0.02099 -0.0044 0.0213 1.0000
12.750 1.3641 0.02912 0.02235 -0.0038 0.0207 1.0000
13.000 1.3749 0.03053 0.02384 -0.0032 0.0200 1.0000
13.250 1.3842 0.03212 0.02548 -0.0027 0.0195 1.0000
13.500 1.3916 0.03392 0.02735 -0.0023 0.0190 1.0000
13.750 1.3962 0.03605 0.02954 -0.0019 0.0186 1.0000
14.000 1.4026 0.03805 0.03163 -0.0016 0.0182 1.0000
14.250 1.4078 0.04022 0.03391 -0.0015 0.0179 1.0000
14.500 1.4115 0.04263 0.03641 -0.0015 0.0176 1.0000
14.750 1.4138 0.04524 0.03912 -0.0016 0.0173 1.0000
15.000 1.4147 0.04810 0.04208 -0.0019 0.0170 1.0000
15.250 1.4146 0.05121 0.04529 -0.0024 0.0167 1.0000
15.500 1.4131 0.05460 0.04877 -0.0032 0.0165 1.0000
15.750 1.4100 0.05832 0.05260 -0.0041 0.0163 1.0000
16.000 1.4054 0.06239 0.05678 -0.0053 0.0161 1.0000
16.250 1.3989 0.06683 0.06132 -0.0068 0.0159 1.0000
16.500 1.3909 0.07161 0.06620 -0.0084 0.0158 1.0000
16.750 1.3811 0.07676 0.07146 -0.0103 0.0156 1.0000
17.000 1.3711 0.08200 0.07681 -0.0123 0.0155 1.0000
17.250 1.3634 0.08693 0.08185 -0.0142 0.0153 1.0000
|
Polar data table (+)
Polar graphs
<< Back to ONERA OA212 AIRFOIL (oa212-il)