RAE 102 AIRFOIL (rae102-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAE 102 AIRFOIL (rae102-il) Reynolds number: 500,000 Max Cl/Cd: 57.25 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rae102-il-500000.txt Download as CSV file: xf-rae102-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: RAE 102 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.500 -0.7060 0.08600 0.08370 -0.0136 1.0000 0.0176
-11.250 -0.7369 0.07371 0.07133 -0.0231 1.0000 0.0172
-11.000 -0.8699 0.05031 0.04717 -0.0345 1.0000 0.0132
-10.750 -0.8874 0.04715 0.04383 -0.0326 1.0000 0.0132
-10.500 -0.8952 0.04523 0.04175 -0.0297 1.0000 0.0131
-10.250 -0.9038 0.04256 0.03887 -0.0267 1.0000 0.0131
-10.000 -0.9103 0.03928 0.03531 -0.0240 1.0000 0.0131
-9.750 -0.9200 0.03466 0.03029 -0.0208 1.0000 0.0132
-9.500 -0.9202 0.03071 0.02597 -0.0181 1.0000 0.0135
-9.250 -0.9103 0.02820 0.02321 -0.0161 1.0000 0.0137
-9.000 -0.8966 0.02626 0.02108 -0.0145 1.0000 0.0141
-8.750 -0.8809 0.02451 0.01913 -0.0130 1.0000 0.0144
-8.500 -0.8634 0.02308 0.01754 -0.0115 1.0000 0.0148
-8.250 -0.8449 0.02177 0.01608 -0.0102 1.0000 0.0154
-8.000 -0.8257 0.02054 0.01470 -0.0089 1.0000 0.0160
-7.750 -0.8061 0.01935 0.01338 -0.0075 1.0000 0.0166
-7.500 -0.7860 0.01841 0.01232 -0.0062 1.0000 0.0174
-7.250 -0.7647 0.01781 0.01161 -0.0050 1.0000 0.0180
-7.000 -0.7503 0.01595 0.00964 -0.0028 1.0000 0.0194
-6.750 -0.7305 0.01530 0.00897 -0.0015 1.0000 0.0208
-6.500 -0.7105 0.01468 0.00831 -0.0001 1.0000 0.0222
-6.250 -0.6905 0.01408 0.00765 0.0014 1.0000 0.0237
-6.000 -0.6695 0.01367 0.00718 0.0027 1.0000 0.0249
-5.750 -0.6544 0.01260 0.00606 0.0049 1.0000 0.0279
-5.500 -0.6340 0.01222 0.00567 0.0063 1.0000 0.0311
-5.250 -0.6134 0.01190 0.00530 0.0077 1.0000 0.0337
-5.000 -0.5961 0.01122 0.00463 0.0096 1.0000 0.0398
-4.750 -0.5765 0.01090 0.00427 0.0111 1.0000 0.0461
-4.500 -0.5479 0.01038 0.00382 0.0106 0.9981 0.0623
-4.250 -0.5114 0.00977 0.00343 0.0083 0.9940 0.1043
-4.000 -0.4781 0.00897 0.00304 0.0065 0.9885 0.1917
-3.750 -0.4445 0.00796 0.00263 0.0044 0.9837 0.3326
-3.500 -0.4138 0.00723 0.00243 0.0033 0.9762 0.4637
-3.250 -0.3766 0.00695 0.00231 0.0012 0.9713 0.5202
-3.000 -0.3425 0.00675 0.00220 -0.0002 0.9631 0.5542
-2.750 -0.3059 0.00660 0.00208 -0.0020 0.9560 0.5824
-2.500 -0.2731 0.00647 0.00199 -0.0030 0.9440 0.6049
-2.250 -0.2410 0.00636 0.00190 -0.0037 0.9303 0.6254
-2.000 -0.2108 0.00627 0.00181 -0.0040 0.9140 0.6443
-1.750 -0.1828 0.00622 0.00174 -0.0038 0.8960 0.6622
-1.500 -0.1562 0.00618 0.00169 -0.0034 0.8769 0.6795
-1.250 -0.1300 0.00615 0.00164 -0.0028 0.8582 0.6959
-1.000 -0.1041 0.00614 0.00160 -0.0022 0.8402 0.7121
-0.750 -0.0782 0.00612 0.00158 -0.0016 0.8224 0.7279
-0.500 -0.0522 0.00611 0.00156 -0.0011 0.8056 0.7434
-0.250 -0.0261 0.00611 0.00155 -0.0005 0.7897 0.7587
0.000 0.0000 0.00611 0.00155 0.0000 0.7741 0.7741
0.250 0.0261 0.00611 0.00155 0.0005 0.7587 0.7897
0.500 0.0522 0.00611 0.00156 0.0010 0.7434 0.8056
0.750 0.0782 0.00612 0.00158 0.0016 0.7278 0.8224
1.000 0.1041 0.00614 0.00160 0.0022 0.7121 0.8402
1.250 0.1300 0.00615 0.00164 0.0028 0.6959 0.8582
1.500 0.1562 0.00618 0.00169 0.0034 0.6795 0.8769
1.750 0.1829 0.00622 0.00174 0.0038 0.6622 0.8960
2.000 0.2108 0.00627 0.00181 0.0040 0.6440 0.9141
2.250 0.2410 0.00636 0.00190 0.0037 0.6252 0.9303
2.500 0.2731 0.00647 0.00199 0.0030 0.6052 0.9441
2.750 0.3059 0.00660 0.00208 0.0020 0.5821 0.9561
3.000 0.3425 0.00675 0.00220 0.0002 0.5548 0.9632
3.250 0.3767 0.00695 0.00230 -0.0012 0.5193 0.9714
3.500 0.4139 0.00723 0.00242 -0.0033 0.4645 0.9764
3.750 0.4446 0.00796 0.00263 -0.0044 0.3308 0.9838
4.000 0.4782 0.00896 0.00303 -0.0065 0.1921 0.9886
4.250 0.5114 0.00977 0.00343 -0.0083 0.1038 0.9941
4.500 0.5482 0.01038 0.00382 -0.0107 0.0622 0.9982
4.750 0.5759 0.01088 0.00426 -0.0110 0.0466 1.0000
5.000 0.5955 0.01122 0.00462 -0.0095 0.0398 1.0000
5.250 0.6128 0.01188 0.00529 -0.0076 0.0337 1.0000
5.500 0.6335 0.01221 0.00566 -0.0062 0.0311 1.0000
5.750 0.6538 0.01261 0.00607 -0.0048 0.0280 1.0000
6.000 0.6690 0.01367 0.00718 -0.0026 0.0249 1.0000
6.250 0.6901 0.01408 0.00766 -0.0013 0.0237 1.0000
6.500 0.7101 0.01468 0.00831 0.0002 0.0222 1.0000
6.750 0.7304 0.01527 0.00894 0.0015 0.0207 1.0000
7.000 0.7500 0.01597 0.00965 0.0029 0.0194 1.0000
7.250 0.7646 0.01779 0.01159 0.0050 0.0180 1.0000
7.500 0.7860 0.01839 0.01230 0.0062 0.0173 1.0000
7.750 0.8061 0.01939 0.01342 0.0075 0.0166 1.0000
8.000 0.8258 0.02055 0.01472 0.0088 0.0160 1.0000
8.250 0.8451 0.02179 0.01610 0.0102 0.0154 1.0000
8.500 0.8637 0.02314 0.01760 0.0115 0.0149 1.0000
8.750 0.8814 0.02447 0.01908 0.0129 0.0144 1.0000
9.000 0.8971 0.02628 0.02111 0.0144 0.0141 1.0000
9.250 0.9106 0.02832 0.02336 0.0161 0.0138 1.0000
9.500 0.9190 0.03110 0.02638 0.0181 0.0134 1.0000
9.750 0.9219 0.03446 0.03007 0.0205 0.0133 1.0000
10.000 0.9104 0.03939 0.03543 0.0238 0.0131 1.0000
10.250 0.9057 0.04247 0.03877 0.0264 0.0131 1.0000
10.500 0.8989 0.04502 0.04153 0.0293 0.0131 1.0000
10.750 0.8826 0.04778 0.04447 0.0326 0.0131 1.0000
11.000 0.8737 0.05008 0.04693 0.0343 0.0133 1.0000
11.250 0.8496 0.05453 0.05157 0.0341 0.0132 1.0000
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Polar data table (+)
Polar graphs
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