RAF 15 AIRFOIL (raf15-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAF 15 AIRFOIL (raf15-il) Reynolds number: 1,000,000 Max Cl/Cd: 92.25 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-raf15-il-1000000.txt Download as CSV file: xf-raf15-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: RAF 15 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
------ -------- --------- --------- -------- -------- --------
-11.500 -0.5568 0.13128 0.12953 0.0080 1.0000 0.0099
-11.250 -0.5540 0.12743 0.12569 0.0069 1.0000 0.0099
-7.250 -0.6986 0.02294 0.01962 -0.0111 1.0000 0.0094
-7.000 -0.6885 0.02031 0.01659 -0.0081 1.0000 0.0097
-6.750 -0.6750 0.01838 0.01430 -0.0054 1.0000 0.0101
-6.500 -0.6636 0.01613 0.01170 -0.0024 1.0000 0.0105
-6.250 -0.6450 0.01527 0.01070 -0.0005 1.0000 0.0108
-6.000 -0.6221 0.01539 0.01086 0.0005 1.0000 0.0114
-5.750 -0.6030 0.01459 0.00990 0.0024 1.0000 0.0117
-5.500 -0.5788 0.01395 0.00914 0.0032 0.9995 0.0122
-5.250 -0.5401 0.01327 0.00831 0.0010 0.9970 0.0128
-5.000 -0.5006 0.01288 0.00781 -0.0015 0.9948 0.0133
-4.750 -0.4693 0.01155 0.00631 -0.0021 0.9920 0.0143
-4.500 -0.4339 0.01134 0.00612 -0.0037 0.9895 0.0153
-4.250 -0.3973 0.01102 0.00576 -0.0055 0.9873 0.0165
-4.000 -0.3598 0.01078 0.00547 -0.0075 0.9858 0.0175
-3.750 -0.3257 0.01010 0.00475 -0.0088 0.9847 0.0192
-3.500 -0.2996 0.00977 0.00442 -0.0082 0.9805 0.0205
-3.250 -0.2674 0.00953 0.00417 -0.0090 0.9779 0.0221
-3.000 -0.2326 0.00939 0.00402 -0.0104 0.9757 0.0233
-2.750 -0.2000 0.00873 0.00332 -0.0113 0.9741 0.0266
-2.500 -0.1743 0.00844 0.00304 -0.0105 0.9685 0.0288
-2.250 -0.1428 0.00820 0.00278 -0.0111 0.9640 0.0306
-2.000 -0.1144 0.00784 0.00240 -0.0110 0.9586 0.0336
-1.500 -0.0635 0.00732 0.00187 -0.0093 0.9374 0.0402
-1.250 -0.0301 0.00687 0.00155 -0.0103 0.9269 0.0695
-1.000 0.0157 0.00631 0.00128 -0.0142 0.9169 0.1472
-0.750 0.1077 0.00604 0.00107 -0.0290 0.8943 0.1882
-0.500 0.1655 0.00615 0.00102 -0.0358 0.8379 0.2055
-0.250 0.1829 0.00647 0.00100 -0.0331 0.7391 0.2141
0.000 0.1990 0.00682 0.00101 -0.0303 0.6544 0.2202
0.250 0.2185 0.00705 0.00103 -0.0283 0.5879 0.2284
0.500 0.2398 0.00727 0.00105 -0.0268 0.5313 0.2378
0.750 0.2623 0.00743 0.00107 -0.0255 0.4856 0.2506
1.000 0.2853 0.00756 0.00110 -0.0244 0.4455 0.2643
1.250 0.3086 0.00767 0.00113 -0.0233 0.4100 0.2806
1.500 0.3317 0.00773 0.00117 -0.0222 0.3822 0.3100
1.750 0.3383 0.00670 0.00118 -0.0178 0.3662 0.7084
2.000 0.4834 0.00701 0.00201 -0.0447 0.3229 0.9845
2.250 0.5331 0.00721 0.00213 -0.0497 0.3057 0.9884
2.500 0.5750 0.00739 0.00223 -0.0528 0.2889 0.9926
2.750 0.6164 0.00753 0.00228 -0.0559 0.2703 0.9956
3.000 0.6561 0.00763 0.00229 -0.0586 0.2492 0.9982
3.250 0.6932 0.00772 0.00231 -0.0607 0.2305 1.0000
3.500 0.7159 0.00787 0.00239 -0.0595 0.2127 1.0000
3.750 0.7383 0.00807 0.00250 -0.0583 0.1924 1.0000
4.000 0.7607 0.00828 0.00263 -0.0571 0.1718 1.0000
4.250 0.7830 0.00850 0.00277 -0.0559 0.1535 1.0000
4.500 0.8053 0.00873 0.00292 -0.0547 0.1369 1.0000
4.750 0.8276 0.00898 0.00311 -0.0535 0.1181 1.0000
5.000 0.8491 0.00931 0.00332 -0.0521 0.0956 1.0000
5.250 0.8696 0.00975 0.00359 -0.0506 0.0654 1.0000
5.500 0.8909 0.01012 0.00388 -0.0492 0.0515 1.0000
5.750 0.9128 0.01044 0.00420 -0.0479 0.0455 1.0000
6.000 0.9355 0.01067 0.00446 -0.0467 0.0437 1.0000
6.250 0.9576 0.01096 0.00476 -0.0455 0.0409 1.0000
6.500 0.9787 0.01136 0.00517 -0.0440 0.0375 1.0000
6.750 1.0000 0.01172 0.00560 -0.0427 0.0359 1.0000
7.000 1.0226 0.01195 0.00587 -0.0415 0.0351 1.0000
7.250 1.0448 0.01222 0.00618 -0.0404 0.0340 1.0000
7.500 1.0668 0.01250 0.00649 -0.0392 0.0325 1.0000
7.750 1.0883 0.01284 0.00685 -0.0379 0.0308 1.0000
8.000 1.1078 0.01336 0.00742 -0.0363 0.0288 1.0000
8.250 1.1273 0.01386 0.00799 -0.0347 0.0273 1.0000
8.500 1.1506 0.01400 0.00815 -0.0338 0.0263 1.0000
8.750 1.1727 0.01425 0.00844 -0.0327 0.0250 1.0000
9.000 1.1944 0.01453 0.00873 -0.0316 0.0232 1.0000
9.250 1.2119 0.01520 0.00943 -0.0297 0.0208 1.0000
9.500 1.2352 0.01531 0.00957 -0.0288 0.0196 1.0000
9.750 1.2566 0.01559 0.00986 -0.0277 0.0179 1.0000
10.000 1.2740 0.01623 0.01051 -0.0258 0.0156 1.0000
10.250 1.2940 0.01662 0.01095 -0.0244 0.0146 1.0000
10.500 1.3130 0.01708 0.01145 -0.0229 0.0135 1.0000
10.750 1.3279 0.01789 0.01231 -0.0206 0.0119 1.0000
11.000 1.3443 0.01854 0.01305 -0.0186 0.0114 1.0000
11.250 1.3615 0.01910 0.01369 -0.0169 0.0107 1.0000
11.500 1.3768 0.01979 0.01446 -0.0148 0.0102 1.0000
11.750 1.3919 0.02048 0.01519 -0.0128 0.0095 1.0000
12.000 1.4003 0.02165 0.01648 -0.0096 0.0088 1.0000
12.250 1.4090 0.02273 0.01771 -0.0066 0.0084 1.0000
12.500 1.4218 0.02348 0.01855 -0.0043 0.0082 1.0000
12.750 1.4315 0.02440 0.01959 -0.0016 0.0080 1.0000
13.000 1.4403 0.02528 0.02056 0.0013 0.0077 1.0000
13.250 1.4412 0.02622 0.02161 0.0056 0.0075 1.0000
13.500 1.4423 0.02716 0.02265 0.0095 0.0073 1.0000
13.750 1.4460 0.02810 0.02367 0.0128 0.0071 1.0000
14.000 1.4432 0.02951 0.02520 0.0166 0.0069 1.0000
14.250 1.4403 0.03107 0.02686 0.0198 0.0067 1.0000
14.500 1.4363 0.03288 0.02880 0.0225 0.0067 1.0000
14.750 1.4163 0.03637 0.03245 0.0248 0.0063 1.0000
15.000 1.4084 0.03941 0.03564 0.0252 0.0064 1.0000
15.250 1.3829 0.04571 0.04215 0.0227 0.0062 1.0000
15.500 1.3732 0.05100 0.04758 0.0194 0.0063 1.0000
15.750 1.3508 0.05907 0.05584 0.0142 0.0063 1.0000
16.000 1.3386 0.06522 0.06212 0.0108 0.0064 1.0000
16.250 1.3111 0.07384 0.07088 0.0061 0.0064 1.0000
16.500 1.2767 0.08342 0.08059 0.0013 0.0064 1.0000
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