RAE 104 AIRFOIL (rae104-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAE 104 AIRFOIL (rae104-il) Reynolds number: 500,000 Max Cl/Cd: 49.7 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rae104-il-500000-n5.txt Download as CSV file: xf-rae104-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAE 104 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
------ -------- --------- --------- -------- -------- --------
-12.250 -0.8703 0.05332 0.05063 -0.0429 1.0000 0.0056
-12.000 -0.8855 0.04915 0.04629 -0.0435 1.0000 0.0058
-11.750 -0.9215 0.04317 0.03999 -0.0423 1.0000 0.0057
-11.500 -0.9329 0.04036 0.03697 -0.0401 1.0000 0.0058
-11.250 -0.9558 0.03651 0.03280 -0.0360 1.0000 0.0057
-11.000 -0.9625 0.03443 0.03051 -0.0322 1.0000 0.0059
-10.750 -0.9696 0.03204 0.02783 -0.0280 1.0000 0.0059
-10.500 -0.9671 0.03003 0.02557 -0.0250 1.0000 0.0061
-10.250 -0.9617 0.02808 0.02336 -0.0222 1.0000 0.0063
-10.000 -0.9534 0.02632 0.02132 -0.0197 1.0000 0.0064
-9.750 -0.9424 0.02481 0.01959 -0.0174 1.0000 0.0066
-9.500 -0.9296 0.02347 0.01805 -0.0152 1.0000 0.0068
-9.250 -0.9145 0.02247 0.01688 -0.0134 1.0000 0.0071
-9.000 -0.8993 0.02144 0.01568 -0.0115 1.0000 0.0072
-8.750 -0.8866 0.02007 0.01416 -0.0092 1.0000 0.0076
-8.500 -0.8709 0.01925 0.01326 -0.0073 1.0000 0.0080
-8.250 -0.8545 0.01856 0.01250 -0.0055 1.0000 0.0083
-8.000 -0.8384 0.01789 0.01174 -0.0036 1.0000 0.0087
-7.750 -0.8121 0.01707 0.01082 -0.0038 0.9973 0.0091
-7.500 -0.7840 0.01627 0.00993 -0.0044 0.9920 0.0097
-7.250 -0.7554 0.01554 0.00910 -0.0051 0.9866 0.0103
-7.000 -0.7255 0.01492 0.00838 -0.0060 0.9818 0.0108
-6.750 -0.6979 0.01404 0.00743 -0.0065 0.9753 0.0117
-6.500 -0.6657 0.01353 0.00687 -0.0079 0.9705 0.0131
-6.250 -0.6352 0.01310 0.00640 -0.0088 0.9632 0.0146
-6.000 -0.6027 0.01266 0.00589 -0.0102 0.9568 0.0158
-5.750 -0.5732 0.01206 0.00525 -0.0109 0.9481 0.0181
-5.500 -0.5418 0.01167 0.00481 -0.0120 0.9400 0.0209
-5.250 -0.5124 0.01133 0.00442 -0.0126 0.9300 0.0239
-5.000 -0.4850 0.01097 0.00405 -0.0127 0.9194 0.0295
-4.750 -0.4582 0.01069 0.00373 -0.0126 0.9091 0.0354
-4.250 -0.4081 0.01011 0.00317 -0.0118 0.8885 0.0612
-4.000 -0.3840 0.00981 0.00292 -0.0112 0.8791 0.0849
-3.500 -0.3381 0.00905 0.00245 -0.0097 0.8622 0.1745
-3.250 -0.3169 0.00853 0.00220 -0.0087 0.8550 0.2531
-3.000 -0.2964 0.00799 0.00198 -0.0076 0.8476 0.3423
-2.750 -0.2776 0.00735 0.00174 -0.0061 0.8412 0.4548
-2.500 -0.2603 0.00673 0.00158 -0.0041 0.8345 0.5725
-2.250 -0.2385 0.00649 0.00156 -0.0028 0.8287 0.6495
-2.000 -0.2134 0.00642 0.00154 -0.0022 0.8233 0.6831
-1.750 -0.1875 0.00638 0.00153 -0.0017 0.8182 0.7047
-1.500 -0.1612 0.00636 0.00151 -0.0014 0.8138 0.7223
-1.250 -0.1347 0.00634 0.00150 -0.0011 0.8087 0.7364
-1.000 -0.1080 0.00633 0.00149 -0.0008 0.8038 0.7482
-0.750 -0.0811 0.00633 0.00148 -0.0006 0.7994 0.7586
-0.500 -0.0541 0.00632 0.00148 -0.0004 0.7942 0.7679
-0.250 -0.0271 0.00631 0.00148 -0.0002 0.7886 0.7756
0.000 0.0000 0.00632 0.00147 0.0000 0.7829 0.7829
0.250 0.0271 0.00631 0.00148 0.0002 0.7756 0.7886
0.500 0.0541 0.00632 0.00148 0.0004 0.7679 0.7942
0.750 0.0811 0.00633 0.00148 0.0006 0.7585 0.7994
1.000 0.1080 0.00633 0.00149 0.0008 0.7482 0.8038
1.250 0.1347 0.00634 0.00150 0.0011 0.7364 0.8088
1.500 0.1612 0.00636 0.00151 0.0014 0.7221 0.8138
1.750 0.1875 0.00638 0.00153 0.0017 0.7046 0.8182
2.000 0.2134 0.00642 0.00154 0.0022 0.6834 0.8233
2.250 0.2385 0.00649 0.00156 0.0028 0.6486 0.8287
2.500 0.2603 0.00672 0.00158 0.0041 0.5731 0.8345
2.750 0.2777 0.00734 0.00174 0.0061 0.4550 0.8412
3.000 0.2964 0.00798 0.00198 0.0076 0.3429 0.8476
3.250 0.3168 0.00854 0.00220 0.0087 0.2518 0.8550
3.500 0.3381 0.00905 0.00245 0.0097 0.1746 0.8622
3.750 0.3607 0.00945 0.00268 0.0105 0.1230 0.8705
4.000 0.3840 0.00981 0.00292 0.0112 0.0844 0.8791
4.250 0.4081 0.01011 0.00317 0.0118 0.0613 0.8886
4.750 0.4582 0.01069 0.00373 0.0126 0.0355 0.9092
5.000 0.4850 0.01097 0.00405 0.0127 0.0295 0.9194
5.250 0.5125 0.01132 0.00441 0.0125 0.0243 0.9300
5.500 0.5418 0.01167 0.00481 0.0120 0.0210 0.9401
5.750 0.5732 0.01205 0.00524 0.0109 0.0182 0.9481
6.000 0.6026 0.01266 0.00589 0.0102 0.0158 0.9569
6.250 0.6351 0.01309 0.00639 0.0089 0.0146 0.9632
6.500 0.6656 0.01353 0.00687 0.0079 0.0131 0.9705
6.750 0.6978 0.01404 0.00742 0.0065 0.0117 0.9753
7.000 0.7255 0.01491 0.00836 0.0060 0.0108 0.9818
7.250 0.7554 0.01553 0.00908 0.0051 0.0103 0.9866
7.500 0.7839 0.01627 0.00992 0.0045 0.0097 0.9920
7.750 0.8121 0.01704 0.01079 0.0038 0.0091 0.9972
8.000 0.8386 0.01787 0.01174 0.0035 0.0087 1.0000
8.250 0.8549 0.01851 0.01245 0.0054 0.0082 1.0000
8.500 0.8708 0.01927 0.01329 0.0073 0.0080 1.0000
8.750 0.8862 0.02014 0.01424 0.0093 0.0077 1.0000
9.000 0.8986 0.02157 0.01583 0.0116 0.0072 1.0000
9.250 0.9140 0.02256 0.01698 0.0134 0.0071 1.0000
9.500 0.9293 0.02353 0.01811 0.0153 0.0069 1.0000
9.750 0.9418 0.02492 0.01972 0.0175 0.0067 1.0000
10.000 0.9532 0.02636 0.02137 0.0197 0.0065 1.0000
10.250 0.9602 0.02831 0.02362 0.0224 0.0064 1.0000
10.500 0.9666 0.03010 0.02565 0.0250 0.0061 1.0000
10.750 0.9680 0.03226 0.02808 0.0282 0.0060 1.0000
11.000 0.9602 0.03470 0.03079 0.0325 0.0059 1.0000
11.250 0.9515 0.03698 0.03332 0.0364 0.0058 1.0000
11.500 0.9409 0.03955 0.03611 0.0396 0.0057 1.0000
11.750 0.9247 0.04287 0.03967 0.0421 0.0056 1.0000
12.000 0.8969 0.04784 0.04492 0.0435 0.0056 1.0000
12.250 0.8783 0.05233 0.04960 0.0430 0.0056 1.0000
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Polar data table (+)
Polar graphs
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