RAE 103 AIRFOIL (rae103-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAE 103 AIRFOIL (rae103-il) Reynolds number: 200,000 Max Cl/Cd: 48.62 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rae103-il-200000.txt Download as CSV file: xf-rae103-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: RAE 103 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.6408 0.09037 0.08684 -0.0178 1.0000 0.0576
-10.000 -0.6566 0.08242 0.07894 -0.0244 1.0000 0.0582
-9.750 -0.6789 0.07551 0.07200 -0.0297 1.0000 0.0581
-9.500 -0.7020 0.07056 0.06699 -0.0318 1.0000 0.0581
-9.250 -0.7232 0.06694 0.06329 -0.0308 1.0000 0.0584
-9.000 -0.7419 0.06338 0.05955 -0.0291 1.0000 0.0600
-8.000 -0.7749 0.04023 0.03471 -0.0179 1.0000 0.0362
-7.750 -0.7704 0.03559 0.02964 -0.0149 1.0000 0.0338
-7.500 -0.7621 0.03122 0.02466 -0.0117 1.0000 0.0321
-7.250 -0.7463 0.02921 0.02241 -0.0099 1.0000 0.0334
-7.000 -0.7293 0.02740 0.02032 -0.0080 1.0000 0.0352
-6.750 -0.7106 0.02515 0.01775 -0.0061 1.0000 0.0358
-6.500 -0.6902 0.02325 0.01556 -0.0045 1.0000 0.0368
-6.250 -0.6690 0.02182 0.01391 -0.0030 1.0000 0.0380
-6.000 -0.6490 0.01981 0.01180 -0.0016 1.0000 0.0408
-5.750 -0.6301 0.01880 0.01082 -0.0001 1.0000 0.0447
-5.500 -0.6113 0.01782 0.00977 0.0017 1.0000 0.0481
-5.250 -0.5951 0.01672 0.00863 0.0039 1.0000 0.0520
-5.000 -0.5801 0.01592 0.00786 0.0061 1.0000 0.0592
-4.750 -0.5667 0.01507 0.00702 0.0087 1.0000 0.0667
-4.500 -0.5534 0.01437 0.00635 0.0112 1.0000 0.0790
-4.250 -0.5413 0.01364 0.00574 0.0138 1.0000 0.1039
-4.000 -0.5336 0.01245 0.00514 0.0167 1.0000 0.1943
-3.750 -0.5282 0.01085 0.00473 0.0196 0.9989 0.4246
-3.500 -0.4969 0.01027 0.00490 0.0187 0.9912 0.6184
-3.250 -0.4591 0.01027 0.00500 0.0168 0.9843 0.6783
-3.000 -0.4235 0.01030 0.00503 0.0154 0.9762 0.7149
-2.750 -0.3842 0.01037 0.00508 0.0133 0.9706 0.7446
-2.500 -0.3507 0.01040 0.00509 0.0123 0.9623 0.7685
-2.250 -0.3117 0.01045 0.00516 0.0104 0.9574 0.7891
-2.000 -0.2773 0.01049 0.00519 0.0094 0.9502 0.8081
-1.750 -0.2392 0.01051 0.00517 0.0076 0.9446 0.8254
-1.500 -0.2016 0.01051 0.00517 0.0060 0.9393 0.8390
-1.250 -0.1675 0.01051 0.00516 0.0050 0.9317 0.8499
-1.000 -0.1285 0.01046 0.00509 0.0029 0.9272 0.8595
-0.750 -0.0995 0.01043 0.00503 0.0027 0.9174 0.8700
-0.500 -0.0619 0.01041 0.00501 0.0010 0.9118 0.8772
-0.250 -0.0327 0.01039 0.00498 0.0008 0.9023 0.8864
0.000 0.0000 0.01039 0.00498 0.0000 0.8946 0.8946
0.250 0.0327 0.01039 0.00498 -0.0008 0.8864 0.9023
0.500 0.0619 0.01041 0.00501 -0.0010 0.8772 0.9119
0.750 0.0995 0.01043 0.00504 -0.0027 0.8700 0.9174
1.000 0.1286 0.01046 0.00509 -0.0029 0.8595 0.9272
1.250 0.1676 0.01051 0.00516 -0.0050 0.8499 0.9317
1.500 0.2016 0.01051 0.00517 -0.0060 0.8389 0.9393
1.750 0.2392 0.01051 0.00518 -0.0077 0.8255 0.9446
2.000 0.2773 0.01049 0.00519 -0.0094 0.8082 0.9502
2.250 0.3117 0.01045 0.00515 -0.0104 0.7889 0.9574
2.500 0.3507 0.01040 0.00509 -0.0123 0.7681 0.9623
2.750 0.3842 0.01036 0.00508 -0.0133 0.7445 0.9706
3.000 0.4235 0.01030 0.00503 -0.0154 0.7150 0.9763
3.250 0.4591 0.01027 0.00500 -0.0168 0.6783 0.9843
3.500 0.4969 0.01026 0.00489 -0.0187 0.6177 0.9912
3.750 0.5280 0.01086 0.00473 -0.0196 0.4212 0.9990
4.000 0.5334 0.01246 0.00515 -0.0167 0.1922 1.0000
4.250 0.5413 0.01363 0.00575 -0.0138 0.1042 1.0000
4.500 0.5534 0.01437 0.00635 -0.0112 0.0791 1.0000
4.750 0.5667 0.01506 0.00702 -0.0087 0.0668 1.0000
5.000 0.5801 0.01592 0.00786 -0.0061 0.0591 1.0000
5.250 0.5950 0.01674 0.00864 -0.0039 0.0519 1.0000
5.500 0.6113 0.01782 0.00977 -0.0017 0.0481 1.0000
5.750 0.6301 0.01881 0.01082 0.0001 0.0447 1.0000
6.000 0.6491 0.01982 0.01181 0.0016 0.0409 1.0000
6.250 0.6690 0.02184 0.01393 0.0030 0.0381 1.0000
6.500 0.6902 0.02326 0.01559 0.0045 0.0367 1.0000
6.750 0.7106 0.02513 0.01772 0.0061 0.0358 1.0000
7.000 0.7293 0.02736 0.02028 0.0080 0.0351 1.0000
7.250 0.7464 0.02921 0.02241 0.0099 0.0335 1.0000
7.500 0.7618 0.03136 0.02484 0.0118 0.0324 1.0000
7.750 0.7704 0.03558 0.02963 0.0150 0.0337 1.0000
8.000 0.7751 0.04020 0.03467 0.0178 0.0363 1.0000
9.000 0.7417 0.06342 0.05960 0.0291 0.0599 1.0000
9.250 0.7234 0.06699 0.06334 0.0307 0.0585 1.0000
9.500 0.7016 0.07067 0.06710 0.0317 0.0581 1.0000
9.750 0.6776 0.07575 0.07224 0.0295 0.0577 1.0000
10.000 0.6564 0.08261 0.07913 0.0241 0.0581 1.0000
10.250 0.6416 0.09038 0.08686 0.0177 0.0578 1.0000
10.500 0.6344 0.09617 0.09261 0.0141 0.0560 1.0000
10.750 0.6312 0.10110 0.09752 0.0115 0.0550 1.0000
11.000 0.6326 0.10501 0.10140 0.0104 0.0535 1.0000
11.250 0.6384 0.10810 0.10450 0.0104 0.0524 1.0000
11.500 0.5122 0.11216 0.10873 0.0073 0.0562 1.0000
11.750 0.5125 0.11612 0.11269 0.0060 0.0545 1.0000
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Polar data table (+)
Polar graphs
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