RAF 25 AIRFOIL (raf25-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: RAF 25 AIRFOIL (raf25-il) Reynolds number: 200,000 Max Cl/Cd: 47.82 at α=2.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-raf25-il-200000-n5.txt Download as CSV file: xf-raf25-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAF 25 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5785 0.09800 0.09440 -0.0052 1.0000 0.0107
-9.250 -0.5835 0.09178 0.08824 -0.0082 1.0000 0.0101
-9.000 -0.5854 0.08714 0.08364 -0.0105 1.0000 0.0101
-8.750 -0.5881 0.08180 0.07836 -0.0134 1.0000 0.0097
-8.500 -0.5899 0.07714 0.07375 -0.0162 1.0000 0.0095
-8.250 -0.5985 0.07162 0.06829 -0.0204 1.0000 0.0095
-8.000 -0.5962 0.06612 0.06277 -0.0252 1.0000 0.0093
-7.750 -0.5968 0.06002 0.05656 -0.0289 1.0000 0.0092
-7.500 -0.5979 0.05362 0.05001 -0.0315 1.0000 0.0091
-7.250 -0.5964 0.04749 0.04364 -0.0327 1.0000 0.0091
-7.000 -0.5927 0.04167 0.03748 -0.0328 1.0000 0.0091
-6.500 -0.5787 0.03090 0.02569 -0.0309 1.0000 0.0099
-6.250 -0.5628 0.02863 0.02312 -0.0299 1.0000 0.0108
-6.000 -0.5459 0.02573 0.01978 -0.0285 1.0000 0.0113
-5.750 -0.5274 0.02292 0.01648 -0.0270 1.0000 0.0118
-5.500 -0.5071 0.02068 0.01384 -0.0255 1.0000 0.0125
-5.250 -0.4857 0.01881 0.01164 -0.0242 1.0000 0.0135
-5.000 -0.4633 0.01767 0.01022 -0.0230 1.0000 0.0157
-4.750 -0.4416 0.01627 0.00862 -0.0217 1.0000 0.0174
-4.500 -0.4204 0.01508 0.00734 -0.0205 1.0000 0.0201
-4.250 -0.3978 0.01445 0.00662 -0.0194 1.0000 0.0252
-4.000 -0.3760 0.01364 0.00571 -0.0183 1.0000 0.0302
-3.750 -0.3533 0.01322 0.00519 -0.0174 1.0000 0.0376
-3.500 -0.3313 0.01265 0.00459 -0.0164 1.0000 0.0447
-3.250 -0.3087 0.01229 0.00415 -0.0154 1.0000 0.0506
-3.000 -0.2865 0.01190 0.00374 -0.0145 1.0000 0.0577
-2.750 -0.2641 0.01159 0.00337 -0.0135 1.0000 0.0652
-2.500 -0.2326 0.01111 0.00298 -0.0145 0.9965 0.0958
-2.250 -0.1987 0.01046 0.00271 -0.0164 0.9907 0.1935
-2.000 -0.1653 0.00988 0.00254 -0.0181 0.9847 0.3136
-1.750 -0.1334 0.00930 0.00242 -0.0193 0.9781 0.4432
-1.500 -0.1042 0.00878 0.00237 -0.0198 0.9702 0.5668
-1.250 -0.0769 0.00820 0.00236 -0.0194 0.9629 0.7090
-1.000 -0.0352 0.00774 0.00244 -0.0215 0.9620 0.8947
-0.750 0.0315 0.00772 0.00241 -0.0298 0.9675 0.9968
-0.500 0.0682 0.00771 0.00233 -0.0318 0.9586 1.0000
-0.250 0.1048 0.00769 0.00225 -0.0338 0.9490 1.0000
0.000 0.1429 0.00765 0.00218 -0.0360 0.9389 1.0000
0.250 0.1820 0.00759 0.00210 -0.0384 0.9243 1.0000
0.500 0.2205 0.00754 0.00203 -0.0405 0.9062 1.0000
0.750 0.2557 0.00752 0.00199 -0.0419 0.8887 1.0000
1.000 0.2879 0.00753 0.00198 -0.0426 0.8626 1.0000
1.250 0.3159 0.00760 0.00196 -0.0422 0.8295 1.0000
1.500 0.3401 0.00772 0.00195 -0.0410 0.7838 1.0000
1.750 0.3622 0.00792 0.00197 -0.0393 0.7237 1.0000
2.000 0.3843 0.00816 0.00201 -0.0378 0.6682 1.0000
2.250 0.4055 0.00848 0.00209 -0.0361 0.6023 1.0000
2.500 0.4253 0.00894 0.00222 -0.0343 0.5178 1.0000
2.750 0.4446 0.00955 0.00245 -0.0326 0.4147 1.0000
3.000 0.4634 0.01036 0.00275 -0.0311 0.2935 1.0000
3.250 0.4806 0.01157 0.00322 -0.0296 0.1283 1.0000
3.500 0.5025 0.01225 0.00367 -0.0286 0.0736 1.0000
3.750 0.5256 0.01279 0.00420 -0.0277 0.0592 1.0000
4.000 0.5487 0.01332 0.00477 -0.0269 0.0459 1.0000
4.250 0.5717 0.01390 0.00543 -0.0260 0.0352 1.0000
4.500 0.5950 0.01446 0.00602 -0.0251 0.0253 1.0000
4.750 0.6166 0.01535 0.00699 -0.0239 0.0210 1.0000
5.000 0.6393 0.01608 0.00785 -0.0228 0.0170 1.0000
5.250 0.6611 0.01710 0.00896 -0.0216 0.0148 1.0000
5.500 0.6809 0.01896 0.01096 -0.0200 0.0132 1.0000
5.750 0.7024 0.02107 0.01331 -0.0187 0.0121 1.0000
6.000 0.7257 0.02230 0.01478 -0.0177 0.0108 1.0000
6.250 0.7475 0.02446 0.01727 -0.0164 0.0101 1.0000
6.500 0.7676 0.02711 0.02033 -0.0148 0.0097 1.0000
6.750 0.7852 0.03025 0.02399 -0.0130 0.0095 1.0000
7.000 0.7997 0.03404 0.02827 -0.0109 0.0093 1.0000
7.250 0.8106 0.03836 0.03308 -0.0087 0.0094 1.0000
7.500 0.8180 0.04302 0.03819 -0.0066 0.0095 1.0000
7.750 0.8234 0.04750 0.04303 -0.0049 0.0093 1.0000
8.000 0.8284 0.05126 0.04706 -0.0035 0.0088 1.0000
8.250 0.8264 0.05627 0.05235 -0.0023 0.0089 1.0000
8.500 0.8200 0.06136 0.05767 -0.0015 0.0091 1.0000
8.750 0.8150 0.06504 0.06149 -0.0012 0.0086 1.0000
9.000 0.7929 0.07092 0.06750 -0.0012 0.0094 1.0000
9.250 0.7772 0.07650 0.07316 -0.0044 0.0092 1.0000
9.500 0.7627 0.08523 0.08188 -0.0122 0.0098 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RAF 25 AIRFOIL (raf25-il)