NASA RC(5)-10 AIRFOIL (rc510-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA RC(5)-10 AIRFOIL (rc510-il) Reynolds number: 100,000 Max Cl/Cd: 43.53 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc510-il-100000-n5.txt Download as CSV file: xf-rc510-il-100000-n5.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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.5370 0.11848 0.11373 0.0066 1.0000 0.0574
-9.750 -0.5399 0.11435 0.10965 0.0019 1.0000 0.0577
-9.500 -0.5409 0.10993 0.10528 -0.0024 1.0000 0.0579
-9.250 -0.5413 0.10522 0.10061 -0.0068 1.0000 0.0579
-9.000 -0.5422 0.10020 0.09563 -0.0118 1.0000 0.0580
-8.750 -0.5433 0.09562 0.09104 -0.0151 1.0000 0.0580
-8.250 -0.5329 0.08110 0.07651 -0.0166 1.0000 0.0366
-8.000 -0.5286 0.07713 0.07254 -0.0182 1.0000 0.0360
-7.750 -0.5249 0.07267 0.06806 -0.0206 1.0000 0.0353
-7.500 -0.5204 0.06797 0.06328 -0.0230 1.0000 0.0345
-7.250 -0.5145 0.06313 0.05833 -0.0252 1.0000 0.0338
-7.000 -0.5068 0.05826 0.05329 -0.0269 1.0000 0.0332
-6.750 -0.4933 0.05307 0.04783 -0.0291 0.9689 0.0326
-6.500 -0.4732 0.04830 0.04271 -0.0312 0.9438 0.0329
-6.250 -0.4537 0.04426 0.03830 -0.0321 0.9263 0.0340
-6.000 -0.4357 0.04028 0.03386 -0.0319 0.9120 0.0346
-5.750 -0.4178 0.03665 0.02974 -0.0311 0.8997 0.0347
-5.500 -0.3988 0.03335 0.02594 -0.0301 0.8890 0.0349
-5.250 -0.3782 0.03044 0.02251 -0.0289 0.8800 0.0353
-5.000 -0.3558 0.02793 0.01948 -0.0278 0.8711 0.0359
-4.750 -0.3317 0.02584 0.01690 -0.0268 0.8636 0.0368
-4.500 -0.3072 0.02443 0.01530 -0.0263 0.8559 0.0386
-4.250 -0.2823 0.02364 0.01438 -0.0258 0.8490 0.0411
-4.000 -0.2562 0.02249 0.01301 -0.0252 0.8418 0.0435
-3.750 -0.2301 0.02129 0.01157 -0.0244 0.8360 0.0458
-3.500 -0.2040 0.02026 0.01049 -0.0240 0.8293 0.0486
-3.250 -0.1785 0.01968 0.00988 -0.0235 0.8236 0.0537
-3.000 -0.1529 0.01902 0.00917 -0.0229 0.8175 0.0604
-2.750 -0.1273 0.01854 0.00864 -0.0224 0.8114 0.0696
-2.500 -0.1026 0.01809 0.00821 -0.0216 0.8066 0.0810
-2.250 -0.0767 0.01770 0.00785 -0.0213 0.8005 0.0957
-2.000 -0.0514 0.01738 0.00750 -0.0207 0.7951 0.1128
-1.750 -0.0265 0.01702 0.00718 -0.0200 0.7903 0.1320
-1.500 -0.0006 0.01670 0.00689 -0.0198 0.7844 0.1558
-1.250 0.0249 0.01631 0.00660 -0.0194 0.7796 0.1902
-1.000 0.0412 0.01430 0.00643 -0.0174 0.7755 0.6454
-0.750 0.1035 0.01398 0.00681 -0.0220 0.7714 0.9280
-0.500 0.1608 0.01415 0.00680 -0.0275 0.7672 0.9765
-0.250 0.2150 0.01417 0.00664 -0.0329 0.7634 1.0000
0.000 0.2396 0.01422 0.00657 -0.0324 0.7558 1.0000
0.250 0.2627 0.01421 0.00642 -0.0313 0.7470 1.0000
0.500 0.2864 0.01420 0.00631 -0.0304 0.7352 1.0000
0.750 0.3099 0.01417 0.00618 -0.0293 0.7238 1.0000
1.000 0.3333 0.01414 0.00603 -0.0282 0.7140 1.0000
1.250 0.3574 0.01413 0.00595 -0.0273 0.7026 1.0000
1.500 0.3815 0.01411 0.00587 -0.0264 0.6903 1.0000
1.750 0.4057 0.01409 0.00579 -0.0254 0.6783 1.0000
2.000 0.4300 0.01407 0.00572 -0.0245 0.6673 1.0000
2.250 0.4545 0.01408 0.00569 -0.0237 0.6558 1.0000
2.500 0.4792 0.01410 0.00570 -0.0229 0.6427 1.0000
2.750 0.5037 0.01411 0.00570 -0.0221 0.6282 1.0000
3.000 0.5282 0.01412 0.00569 -0.0212 0.6120 1.0000
3.250 0.5525 0.01413 0.00567 -0.0203 0.5930 1.0000
3.500 0.5768 0.01418 0.00571 -0.0194 0.5685 1.0000
3.750 0.6006 0.01424 0.00571 -0.0184 0.5365 1.0000
4.000 0.6237 0.01438 0.00569 -0.0173 0.4855 1.0000
4.250 0.6442 0.01480 0.00565 -0.0158 0.4039 1.0000
4.500 0.6638 0.01555 0.00597 -0.0146 0.3443 1.0000
4.750 0.6845 0.01624 0.00641 -0.0137 0.3103 1.0000
5.000 0.7061 0.01684 0.00687 -0.0128 0.2869 1.0000
5.250 0.7278 0.01740 0.00734 -0.0120 0.2697 1.0000
5.500 0.7498 0.01794 0.00783 -0.0112 0.2559 1.0000
5.750 0.7718 0.01847 0.00832 -0.0103 0.2439 1.0000
6.000 0.7936 0.01904 0.00881 -0.0095 0.2336 1.0000
6.250 0.8161 0.01955 0.00933 -0.0087 0.2238 1.0000
6.500 0.8382 0.02011 0.00989 -0.0080 0.2154 1.0000
6.750 0.8604 0.02066 0.01045 -0.0072 0.2070 1.0000
7.000 0.8827 0.02124 0.01105 -0.0065 0.1994 1.0000
7.250 0.9048 0.02183 0.01165 -0.0057 0.1922 1.0000
7.500 0.9269 0.02245 0.01233 -0.0050 0.1853 1.0000
7.750 0.9489 0.02305 0.01299 -0.0043 0.1783 1.0000
8.000 0.9706 0.02373 0.01369 -0.0036 0.1723 1.0000
8.250 0.9924 0.02436 0.01444 -0.0029 0.1654 1.0000
8.500 1.0134 0.02508 0.01515 -0.0022 0.1599 1.0000
8.750 1.0349 0.02576 0.01604 -0.0015 0.1533 1.0000
9.000 1.0552 0.02647 0.01678 -0.0007 0.1477 1.0000
9.250 1.0756 0.02724 0.01770 0.0001 0.1416 1.0000
9.500 1.0952 0.02798 0.01857 0.0009 0.1356 1.0000
9.750 1.1141 0.02880 0.01946 0.0017 0.1303 1.0000
10.000 1.1323 0.02960 0.02046 0.0027 0.1238 1.0000
10.250 1.1492 0.03041 0.02127 0.0036 0.1188 1.0000
10.500 1.1659 0.03136 0.02251 0.0046 0.1124 1.0000
10.750 1.1807 0.03220 0.02343 0.0057 0.1071 1.0000
11.000 1.1948 0.03328 0.02469 0.0069 0.1016 1.0000
11.250 1.2069 0.03427 0.02584 0.0082 0.0960 1.0000
11.500 1.2162 0.03539 0.02701 0.0097 0.0916 1.0000
11.750 1.2241 0.03668 0.02855 0.0112 0.0861 1.0000
12.000 1.2293 0.03796 0.02991 0.0126 0.0818 1.0000
12.250 1.2348 0.03964 0.03179 0.0137 0.0771 1.0000
12.500 1.2387 0.04138 0.03368 0.0145 0.0725 1.0000
12.750 1.2399 0.04332 0.03561 0.0152 0.0691 1.0000
13.000 1.2416 0.04569 0.03829 0.0157 0.0647 1.0000
13.250 1.2410 0.04813 0.04087 0.0159 0.0611 1.0000
13.500 1.2377 0.05086 0.04359 0.0159 0.0584 1.0000
13.750 1.2334 0.05417 0.04721 0.0158 0.0553 1.0000
14.000 1.2271 0.05772 0.05095 0.0152 0.0526 1.0000
14.250 1.2189 0.06158 0.05492 0.0141 0.0505 1.0000
14.500 1.2088 0.06581 0.05920 0.0127 0.0488 1.0000
14.750 1.1950 0.07105 0.06468 0.0109 0.0473 1.0000
15.000 1.1783 0.07701 0.07087 0.0083 0.0460 1.0000
15.250 1.1592 0.08381 0.07788 0.0050 0.0452 1.0000
15.500 1.1371 0.09171 0.08598 0.0008 0.0446 1.0000
15.750 1.1104 0.10112 0.09557 -0.0043 0.0445 1.0000
16.000 1.0751 0.11290 0.10755 -0.0108 0.0451 1.0000
16.250 1.0204 0.12986 0.12467 -0.0199 0.0466 1.0000
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