ONERA NACA CAMBRE AIRFOIL (ncambre-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: ONERA NACA CAMBRE AIRFOIL (ncambre-il) Reynolds number: 1,000,000 Max Cl/Cd: 90.41 at α=8° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ncambre-il-1000000.txt Download as CSV file: xf-ncambre-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.6741 0.07233 0.07053 -0.0282 1.0000 0.0186
-10.500 -0.7052 0.06413 0.06218 -0.0343 1.0000 0.0186
-10.250 -0.7312 0.05903 0.05695 -0.0353 1.0000 0.0187
-10.000 -0.7497 0.05555 0.05336 -0.0334 1.0000 0.0188
-9.750 -0.7527 0.05285 0.05057 -0.0322 1.0000 0.0188
-9.500 -0.7518 0.05035 0.04797 -0.0309 1.0000 0.0189
-9.250 -0.7475 0.04809 0.04561 -0.0297 1.0000 0.0191
-9.000 -0.7400 0.04617 0.04361 -0.0284 1.0000 0.0192
-8.750 -0.7313 0.04434 0.04169 -0.0270 1.0000 0.0195
-8.500 -0.7657 0.03281 0.02935 -0.0196 1.0000 0.0178
-8.250 -0.7671 0.02888 0.02508 -0.0156 1.0000 0.0175
-8.000 -0.7671 0.02475 0.02051 -0.0112 1.0000 0.0172
-7.750 -0.7593 0.02187 0.01729 -0.0079 1.0000 0.0173
-7.500 -0.7401 0.01986 0.01502 -0.0066 0.9994 0.0175
-7.250 -0.7090 0.01845 0.01341 -0.0075 0.9979 0.0179
-7.000 -0.6756 0.01778 0.01260 -0.0089 0.9962 0.0183
-6.750 -0.6442 0.01607 0.01071 -0.0100 0.9948 0.0187
-6.500 -0.6130 0.01476 0.00934 -0.0110 0.9931 0.0191
-6.250 -0.5813 0.01408 0.00864 -0.0120 0.9905 0.0194
-6.000 -0.5486 0.01350 0.00803 -0.0132 0.9883 0.0198
-5.750 -0.5150 0.01292 0.00744 -0.0146 0.9864 0.0202
-5.500 -0.4802 0.01236 0.00685 -0.0162 0.9846 0.0207
-5.250 -0.4486 0.01183 0.00630 -0.0170 0.9803 0.0212
-5.000 -0.4182 0.01137 0.00581 -0.0176 0.9745 0.0217
-4.750 -0.3879 0.01102 0.00543 -0.0180 0.9685 0.0222
-4.500 -0.3631 0.01038 0.00476 -0.0174 0.9589 0.0230
-4.250 -0.3378 0.00997 0.00434 -0.0167 0.9487 0.0237
-4.000 -0.3120 0.00967 0.00403 -0.0162 0.9378 0.0244
-3.750 -0.2866 0.00941 0.00374 -0.0155 0.9241 0.0252
-3.500 -0.2613 0.00918 0.00345 -0.0147 0.9073 0.0260
-3.250 -0.2354 0.00901 0.00321 -0.0141 0.8894 0.0268
-3.000 -0.2100 0.00869 0.00282 -0.0134 0.8717 0.0281
-2.750 -0.1837 0.00852 0.00260 -0.0130 0.8523 0.0297
-2.500 -0.1573 0.00842 0.00242 -0.0125 0.8305 0.0315
-2.250 -0.1307 0.00830 0.00221 -0.0121 0.8080 0.0335
-2.000 -0.1042 0.00817 0.00202 -0.0117 0.7842 0.0369
-1.750 -0.0773 0.00809 0.00185 -0.0113 0.7586 0.0414
-1.250 -0.0262 0.00747 0.00151 -0.0104 0.7098 0.1684
-1.000 -0.0068 0.00626 0.00125 -0.0094 0.6893 0.4648
-0.750 0.0180 0.00593 0.00120 -0.0088 0.6692 0.5627
-0.500 0.0433 0.00571 0.00117 -0.0083 0.6491 0.6397
-0.250 0.0689 0.00556 0.00116 -0.0078 0.6258 0.7021
0.000 0.0942 0.00548 0.00117 -0.0071 0.5997 0.7565
0.250 0.1204 0.00548 0.00119 -0.0066 0.5748 0.7954
0.500 0.1459 0.00547 0.00123 -0.0059 0.5504 0.8345
0.750 0.1719 0.00552 0.00128 -0.0053 0.5252 0.8640
1.000 0.1983 0.00559 0.00132 -0.0048 0.5046 0.8858
1.250 0.2250 0.00568 0.00137 -0.0044 0.4863 0.9016
1.500 0.2515 0.00578 0.00144 -0.0039 0.4688 0.9174
1.750 0.2779 0.00589 0.00151 -0.0034 0.4497 0.9317
2.000 0.3043 0.00604 0.00158 -0.0030 0.4263 0.9434
2.250 0.3307 0.00620 0.00165 -0.0026 0.4009 0.9538
2.500 0.3587 0.00640 0.00174 -0.0025 0.3757 0.9613
2.750 0.3854 0.00658 0.00183 -0.0022 0.3541 0.9699
3.000 0.4172 0.00676 0.00192 -0.0030 0.3346 0.9741
3.250 0.4491 0.00695 0.00203 -0.0040 0.3149 0.9782
3.500 0.4786 0.00714 0.00213 -0.0043 0.2949 0.9833
3.750 0.5138 0.00735 0.00225 -0.0061 0.2759 0.9848
4.000 0.5494 0.00754 0.00237 -0.0079 0.2600 0.9863
4.250 0.5847 0.00774 0.00250 -0.0096 0.2449 0.9882
4.500 0.6193 0.00793 0.00263 -0.0112 0.2304 0.9906
4.750 0.6528 0.00814 0.00278 -0.0125 0.2164 0.9932
5.000 0.6876 0.00836 0.00292 -0.0142 0.2031 0.9949
5.250 0.7229 0.00857 0.00307 -0.0160 0.1912 0.9964
5.500 0.7581 0.00875 0.00323 -0.0177 0.1806 0.9980
5.750 0.7932 0.00896 0.00339 -0.0195 0.1695 0.9996
6.000 0.8192 0.00917 0.00356 -0.0193 0.1591 1.0000
6.250 0.8417 0.00937 0.00372 -0.0183 0.1497 1.0000
6.500 0.8642 0.00958 0.00390 -0.0174 0.1410 1.0000
6.750 0.8862 0.00984 0.00411 -0.0163 0.1316 1.0000
7.000 0.9085 0.01007 0.00431 -0.0153 0.1237 1.0000
7.250 0.9308 0.01031 0.00452 -0.0143 0.1155 1.0000
7.500 0.9529 0.01057 0.00475 -0.0133 0.1087 1.0000
7.750 0.9755 0.01080 0.00498 -0.0124 0.1039 1.0000
8.000 0.9981 0.01104 0.00522 -0.0115 0.0991 1.0000
8.250 1.0202 0.01135 0.00550 -0.0105 0.0933 1.0000
8.500 1.0435 0.01156 0.00573 -0.0097 0.0905 1.0000
8.750 1.0663 0.01183 0.00600 -0.0089 0.0865 1.0000
9.000 1.0887 0.01218 0.00634 -0.0080 0.0816 1.0000
9.250 1.1129 0.01238 0.00658 -0.0075 0.0792 1.0000
9.500 1.1363 0.01269 0.00688 -0.0068 0.0747 1.0000
9.750 1.1596 0.01303 0.00721 -0.0062 0.0696 1.0000
10.000 1.1828 0.01338 0.00754 -0.0056 0.0635 1.0000
10.250 1.2052 0.01381 0.00794 -0.0049 0.0547 1.0000
10.500 1.2244 0.01456 0.00856 -0.0038 0.0389 1.0000
10.750 1.2431 0.01535 0.00929 -0.0027 0.0299 1.0000
11.000 1.2623 0.01607 0.01001 -0.0016 0.0256 1.0000
11.250 1.2831 0.01662 0.01059 -0.0007 0.0235 1.0000
11.500 1.3015 0.01737 0.01137 0.0005 0.0214 1.0000
11.750 1.3219 0.01791 0.01196 0.0014 0.0205 1.0000
12.000 1.3413 0.01850 0.01260 0.0023 0.0196 1.0000
12.250 1.3591 0.01919 0.01333 0.0035 0.0187 1.0000
12.500 1.3736 0.02009 0.01430 0.0051 0.0177 1.0000
12.750 1.3894 0.02084 0.01511 0.0065 0.0172 1.0000
13.000 1.4041 0.02149 0.01583 0.0081 0.0167 1.0000
13.250 1.4162 0.02223 0.01665 0.0100 0.0162 1.0000
13.500 1.4272 0.02309 0.01757 0.0119 0.0158 1.0000
13.750 1.4365 0.02408 0.01863 0.0138 0.0154 1.0000
14.000 1.4433 0.02531 0.01994 0.0156 0.0150 1.0000
14.250 1.4456 0.02694 0.02166 0.0175 0.0146 1.0000
14.500 1.4417 0.02923 0.02407 0.0192 0.0142 1.0000
14.750 1.4477 0.03089 0.02583 0.0199 0.0140 1.0000
15.000 1.4529 0.03276 0.02780 0.0204 0.0138 1.0000
15.250 1.4560 0.03498 0.03012 0.0205 0.0136 1.0000
15.500 1.4576 0.03754 0.03279 0.0203 0.0134 1.0000
15.750 1.4566 0.04058 0.03594 0.0196 0.0132 1.0000
16.000 1.4536 0.04407 0.03954 0.0185 0.0130 1.0000
16.250 1.4475 0.04814 0.04373 0.0169 0.0128 1.0000
16.500 1.4384 0.05281 0.04853 0.0149 0.0127 1.0000
16.750 1.4254 0.05820 0.05405 0.0125 0.0126 1.0000
17.000 1.4089 0.06418 0.06016 0.0098 0.0125 1.0000
17.250 1.3883 0.07090 0.06702 0.0067 0.0124 1.0000
17.500 1.3656 0.07806 0.07432 0.0035 0.0124 1.0000
17.750 1.3408 0.08574 0.08213 -0.0001 0.0124 1.0000
18.000 1.3155 0.09357 0.09008 -0.0037 0.0124 1.0000
|
Polar data table (+)
Polar graphs
<< Back to ONERA NACA CAMBRE AIRFOIL (ncambre-il)