NASA RC(5)-10 AIRFOIL (rc510-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA RC(5)-10 AIRFOIL (rc510-il) Reynolds number: 500,000 Max Cl/Cd: 76.31 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc510-il-500000.txt Download as CSV file: xf-rc510-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NASA RC(5)-10 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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.5433 0.08642 0.08446 0.0001 0.9638 0.0247
-8.250 -0.5489 0.08205 0.07994 -0.0032 0.8911 0.0251
-8.000 -0.5540 0.07690 0.07466 -0.0082 0.8692 0.0254
-7.750 -0.5530 0.07148 0.06911 -0.0126 0.8563 0.0262
-7.500 -0.5484 0.06281 0.06007 -0.0191 0.8463 0.0279
-7.250 -0.5516 0.05477 0.05176 -0.0205 0.8382 0.0282
-7.000 -0.5424 0.05156 0.04845 -0.0204 0.8310 0.0287
-6.750 -0.5279 0.04929 0.04609 -0.0203 0.8238 0.0291
-6.500 -0.5118 0.04705 0.04372 -0.0201 0.8175 0.0297
-6.250 -0.4944 0.04447 0.04100 -0.0201 0.8110 0.0306
-6.000 -0.4766 0.04124 0.03754 -0.0197 0.8056 0.0323
-5.750 -0.4837 0.02185 0.01610 -0.0143 0.8020 0.0228
-5.500 -0.4600 0.01990 0.01383 -0.0135 0.7964 0.0227
-5.250 -0.4361 0.01795 0.01159 -0.0127 0.7914 0.0229
-5.000 -0.4109 0.01654 0.01000 -0.0122 0.7865 0.0234
-4.750 -0.3848 0.01562 0.00898 -0.0118 0.7810 0.0239
-4.500 -0.3585 0.01488 0.00814 -0.0114 0.7761 0.0245
-4.250 -0.3319 0.01421 0.00738 -0.0110 0.7716 0.0252
-4.000 -0.3051 0.01361 0.00671 -0.0107 0.7665 0.0262
-3.750 -0.2781 0.01319 0.00621 -0.0103 0.7619 0.0274
-3.500 -0.2523 0.01251 0.00543 -0.0098 0.7577 0.0286
-3.250 -0.2262 0.01189 0.00482 -0.0094 0.7526 0.0303
-3.000 -0.1993 0.01153 0.00443 -0.0090 0.7478 0.0324
-2.750 -0.1723 0.01128 0.00411 -0.0087 0.7439 0.0347
-2.500 -0.1458 0.01080 0.00367 -0.0083 0.7394 0.0401
-2.250 -0.1187 0.01049 0.00338 -0.0081 0.7347 0.0488
-2.000 -0.0915 0.01025 0.00316 -0.0078 0.7306 0.0618
-1.750 -0.0640 0.01008 0.00301 -0.0077 0.7264 0.0763
-1.500 -0.0362 0.00991 0.00289 -0.0076 0.7204 0.0915
-1.250 -0.0089 0.00974 0.00273 -0.0073 0.7138 0.1080
-1.000 0.0186 0.00954 0.00260 -0.0072 0.7063 0.1271
-0.750 0.0460 0.00935 0.00244 -0.0070 0.6997 0.1503
-0.500 0.0728 0.00903 0.00230 -0.0068 0.6936 0.2002
-0.250 0.0856 0.00713 0.00211 -0.0043 0.6876 0.6972
0.000 0.1083 0.00653 0.00221 -0.0022 0.6820 0.9133
0.250 0.1569 0.00663 0.00233 -0.0061 0.6744 0.9617
0.500 0.2023 0.00677 0.00237 -0.0095 0.6670 0.9776
0.750 0.2496 0.00685 0.00240 -0.0135 0.6580 0.9885
1.000 0.2966 0.00690 0.00236 -0.0175 0.6491 0.9964
1.250 0.3354 0.00690 0.00231 -0.0198 0.6395 1.0000
1.500 0.3625 0.00689 0.00226 -0.0197 0.6314 1.0000
1.750 0.3895 0.00689 0.00220 -0.0195 0.6221 1.0000
2.000 0.4165 0.00688 0.00217 -0.0193 0.6108 1.0000
2.250 0.4434 0.00690 0.00214 -0.0191 0.5988 1.0000
2.500 0.4701 0.00693 0.00212 -0.0189 0.5856 1.0000
2.750 0.4967 0.00697 0.00212 -0.0186 0.5682 1.0000
3.000 0.5232 0.00706 0.00212 -0.0183 0.5442 1.0000
3.250 0.5494 0.00720 0.00214 -0.0180 0.5063 1.0000
3.500 0.5743 0.00772 0.00225 -0.0178 0.4108 1.0000
3.750 0.5982 0.00852 0.00258 -0.0176 0.3178 1.0000
4.000 0.6227 0.00898 0.00283 -0.0173 0.2782 1.0000
4.250 0.6475 0.00931 0.00304 -0.0169 0.2554 1.0000
4.500 0.6723 0.00960 0.00326 -0.0165 0.2391 1.0000
4.750 0.6970 0.00988 0.00347 -0.0160 0.2259 1.0000
5.250 0.7463 0.01040 0.00391 -0.0151 0.2054 1.0000
5.500 0.7706 0.01069 0.00416 -0.0145 0.1970 1.0000
5.750 0.7952 0.01093 0.00438 -0.0140 0.1891 1.0000
6.000 0.8193 0.01123 0.00466 -0.0135 0.1820 1.0000
6.250 0.8438 0.01146 0.00491 -0.0130 0.1755 1.0000
6.500 0.8674 0.01181 0.00522 -0.0124 0.1687 1.0000
6.750 0.8921 0.01202 0.00546 -0.0119 0.1631 1.0000
7.000 0.9156 0.01236 0.00577 -0.0113 0.1572 1.0000
7.250 0.9396 0.01264 0.00609 -0.0107 0.1518 1.0000
7.500 0.9636 0.01290 0.00637 -0.0101 0.1462 1.0000
7.750 0.9861 0.01333 0.00676 -0.0094 0.1403 1.0000
8.000 1.0107 0.01353 0.00703 -0.0089 0.1352 1.0000
8.250 1.0337 0.01389 0.00737 -0.0083 0.1295 1.0000
8.500 1.0569 0.01423 0.00776 -0.0077 0.1244 1.0000
8.750 1.0806 0.01452 0.00807 -0.0072 0.1190 1.0000
9.000 1.1024 0.01501 0.00854 -0.0065 0.1131 1.0000
9.250 1.1266 0.01524 0.00885 -0.0060 0.1080 1.0000
9.500 1.1484 0.01571 0.00928 -0.0054 0.1019 1.0000
9.750 1.1717 0.01602 0.00967 -0.0049 0.0965 1.0000
10.000 1.1935 0.01648 0.01011 -0.0043 0.0898 1.0000
10.250 1.2160 0.01686 0.01054 -0.0037 0.0839 1.0000
10.500 1.2365 0.01742 0.01107 -0.0030 0.0771 1.0000
10.750 1.2583 0.01785 0.01156 -0.0024 0.0715 1.0000
11.000 1.2773 0.01852 0.01221 -0.0016 0.0649 1.0000
11.250 1.2977 0.01904 0.01278 -0.0009 0.0591 1.0000
11.500 1.3157 0.01975 0.01350 0.0000 0.0531 1.0000
11.750 1.3326 0.02053 0.01427 0.0010 0.0472 1.0000
12.000 1.3489 0.02131 0.01510 0.0021 0.0420 1.0000
12.250 1.3599 0.02236 0.01616 0.0037 0.0368 1.0000
12.500 1.3717 0.02329 0.01714 0.0052 0.0328 1.0000
12.750 1.3795 0.02461 0.01850 0.0067 0.0289 1.0000
13.000 1.3881 0.02597 0.01990 0.0078 0.0256 1.0000
13.250 1.3940 0.02765 0.02165 0.0089 0.0229 1.0000
13.500 1.4011 0.02928 0.02335 0.0096 0.0207 1.0000
13.750 1.4006 0.03168 0.02583 0.0105 0.0189 1.0000
14.000 1.4070 0.03348 0.02775 0.0110 0.0176 1.0000
14.250 1.4103 0.03564 0.03001 0.0114 0.0165 1.0000
14.500 1.4085 0.03837 0.03283 0.0117 0.0157 1.0000
14.750 1.3987 0.04206 0.03663 0.0116 0.0150 1.0000
15.000 1.3970 0.04500 0.03971 0.0115 0.0145 1.0000
15.250 1.3938 0.04821 0.04306 0.0111 0.0141 1.0000
15.500 1.3878 0.05186 0.04684 0.0104 0.0137 1.0000
15.750 1.3797 0.05594 0.05105 0.0094 0.0134 1.0000
16.000 1.3692 0.06052 0.05577 0.0080 0.0131 1.0000
16.250 1.3572 0.06568 0.06106 0.0060 0.0129 1.0000
16.500 1.3428 0.07150 0.06702 0.0036 0.0128 1.0000
16.750 1.3260 0.07804 0.07371 0.0006 0.0126 1.0000
17.000 1.3074 0.08521 0.08103 -0.0028 0.0126 1.0000
17.250 1.2876 0.09275 0.08871 -0.0064 0.0125 1.0000
17.500 1.2676 0.10042 0.09651 -0.0101 0.0125 1.0000
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