NASA RC(4)-10 AIRFOIL (rc410-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA RC(4)-10 AIRFOIL (rc410-il) Reynolds number: 200,000 Max Cl/Cd: 47.87 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc410-il-200000.txt Download as CSV file: xf-rc410-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NASA RC(4)-10 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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.5259 0.09192 0.08870 0.0025 1.0000 0.0654
-8.250 -0.5278 0.08713 0.08395 -0.0025 1.0000 0.0674
-8.000 -0.5563 0.07840 0.07500 -0.0203 1.0000 0.0699
-7.750 -0.5534 0.07206 0.06862 -0.0225 1.0000 0.0705
-7.500 -0.5387 0.06822 0.06486 -0.0225 1.0000 0.0712
-7.250 -0.5234 0.06506 0.06173 -0.0232 1.0000 0.0721
-7.000 -0.5082 0.06181 0.05847 -0.0246 1.0000 0.0733
-6.750 -0.5153 0.04262 0.03826 -0.0313 0.9716 0.0493
-6.500 -0.4989 0.03510 0.03014 -0.0326 0.9542 0.0479
-6.250 -0.4868 0.02783 0.02194 -0.0315 0.9395 0.0473
-6.000 -0.4679 0.02431 0.01780 -0.0301 0.9269 0.0479
-5.750 -0.4458 0.02218 0.01520 -0.0289 0.9163 0.0492
-5.500 -0.4228 0.02073 0.01332 -0.0276 0.9071 0.0508
-5.250 -0.3990 0.01955 0.01212 -0.0268 0.8976 0.0533
-5.000 -0.3745 0.01913 0.01161 -0.0259 0.8896 0.0568
-4.750 -0.3492 0.01814 0.01043 -0.0252 0.8817 0.0611
-4.500 -0.3243 0.01778 0.01009 -0.0243 0.8752 0.0660
-4.250 -0.2980 0.01721 0.00946 -0.0239 0.8679 0.0733
-4.000 -0.2726 0.01707 0.00921 -0.0231 0.8616 0.0830
-3.750 -0.2458 0.01714 0.00927 -0.0228 0.8548 0.0927
-3.500 -0.2198 0.01699 0.00915 -0.0223 0.8485 0.1015
-3.250 -0.1940 0.01685 0.00889 -0.0217 0.8436 0.1110
-3.000 -0.1663 0.01689 0.00893 -0.0217 0.8374 0.1203
-2.750 -0.1400 0.01655 0.00858 -0.0213 0.8320 0.1285
-2.500 -0.1136 0.01645 0.00839 -0.0208 0.8272 0.1365
-2.250 -0.0862 0.01596 0.00793 -0.0207 0.8204 0.1434
-2.000 -0.0609 0.01573 0.00764 -0.0198 0.8135 0.1507
-1.750 -0.0348 0.01533 0.00722 -0.0192 0.8045 0.1571
-1.500 -0.0102 0.01499 0.00688 -0.0180 0.7968 0.1644
-1.250 0.0165 0.01473 0.00660 -0.0176 0.7879 0.1712
-1.000 0.0410 0.01431 0.00623 -0.0165 0.7810 0.1793
-0.750 0.0675 0.01405 0.00599 -0.0161 0.7728 0.1881
-0.500 0.0925 0.01371 0.00570 -0.0152 0.7658 0.1978
-0.250 0.1181 0.01340 0.00548 -0.0145 0.7578 0.2105
0.000 0.1432 0.01308 0.00521 -0.0136 0.7500 0.2292
0.250 0.1657 0.01228 0.00497 -0.0127 0.7417 0.3547
0.500 0.2508 0.01071 0.00521 -0.0219 0.7318 0.9905
0.750 0.2948 0.01061 0.00501 -0.0250 0.7197 1.0000
1.000 0.3198 0.01053 0.00484 -0.0241 0.7090 1.0000
1.250 0.3443 0.01042 0.00462 -0.0230 0.6980 1.0000
1.500 0.3693 0.01031 0.00445 -0.0222 0.6839 1.0000
1.750 0.3944 0.01021 0.00429 -0.0213 0.6686 1.0000
2.000 0.4193 0.01011 0.00412 -0.0203 0.6511 1.0000
2.250 0.4445 0.01003 0.00399 -0.0195 0.6273 1.0000
2.500 0.4690 0.00998 0.00380 -0.0184 0.5892 1.0000
2.750 0.4911 0.01026 0.00354 -0.0170 0.4581 1.0000
3.000 0.5107 0.01159 0.00392 -0.0162 0.3139 1.0000
3.250 0.5338 0.01224 0.00430 -0.0156 0.2840 1.0000
3.500 0.5573 0.01274 0.00464 -0.0149 0.2667 1.0000
3.750 0.5810 0.01321 0.00497 -0.0143 0.2534 1.0000
4.000 0.6050 0.01358 0.00529 -0.0136 0.2425 1.0000
4.250 0.6284 0.01409 0.00567 -0.0129 0.2338 1.0000
4.500 0.6525 0.01445 0.00601 -0.0122 0.2259 1.0000
4.750 0.6756 0.01502 0.00645 -0.0114 0.2191 1.0000
5.000 0.7000 0.01535 0.00681 -0.0107 0.2123 1.0000
5.250 0.7236 0.01581 0.00719 -0.0100 0.2062 1.0000
5.500 0.7473 0.01634 0.00769 -0.0093 0.2009 1.0000
5.750 0.7714 0.01674 0.00812 -0.0086 0.1956 1.0000
6.000 0.7951 0.01723 0.00855 -0.0080 0.1906 1.0000
6.250 0.8188 0.01785 0.00915 -0.0073 0.1858 1.0000
6.500 0.8428 0.01825 0.00962 -0.0066 0.1810 1.0000
6.750 0.8667 0.01874 0.01011 -0.0060 0.1767 1.0000
7.000 0.8906 0.01959 0.01084 -0.0055 0.1724 1.0000
7.250 0.9143 0.01999 0.01139 -0.0048 0.1685 1.0000
7.500 0.9381 0.02046 0.01194 -0.0041 0.1641 1.0000
7.750 0.9620 0.02101 0.01247 -0.0036 0.1604 1.0000
8.000 0.9860 0.02204 0.01343 -0.0032 0.1567 1.0000
8.250 1.0090 0.02247 0.01405 -0.0025 0.1531 1.0000
8.500 1.0321 0.02300 0.01468 -0.0018 0.1491 1.0000
8.750 1.0557 0.02358 0.01527 -0.0013 0.1457 1.0000
9.000 1.0795 0.02479 0.01640 -0.0010 0.1422 1.0000
9.250 1.1011 0.02528 0.01715 -0.0002 0.1390 1.0000
9.500 1.1230 0.02591 0.01791 0.0006 0.1352 1.0000
9.750 1.1455 0.02650 0.01854 0.0011 0.1320 1.0000
10.000 1.1688 0.02744 0.01940 0.0015 0.1288 1.0000
10.250 1.1884 0.02838 0.02059 0.0024 0.1257 1.0000
10.500 1.2081 0.02914 0.02155 0.0032 0.1222 1.0000
10.750 1.2288 0.02971 0.02218 0.0040 0.1189 1.0000
11.000 1.2510 0.03037 0.02279 0.0044 0.1160 1.0000
11.250 1.2685 0.03171 0.02432 0.0053 0.1130 1.0000
11.500 1.2844 0.03261 0.02547 0.0065 0.1097 1.0000
11.750 1.3019 0.03325 0.02622 0.0075 0.1065 1.0000
12.000 1.3222 0.03374 0.02665 0.0081 0.1038 1.0000
12.250 1.3372 0.03522 0.02826 0.0091 0.1009 1.0000
12.500 1.3467 0.03632 0.02965 0.0107 0.0979 1.0000
12.750 1.3584 0.03715 0.03062 0.0121 0.0949 1.0000
13.000 1.3748 0.03747 0.03091 0.0131 0.0922 1.0000
13.250 1.3876 0.03887 0.03231 0.0141 0.0894 1.0000
13.500 1.3843 0.04027 0.03401 0.0167 0.0873 1.0000
13.750 1.3830 0.04172 0.03567 0.0186 0.0849 1.0000
14.000 1.3892 0.04263 0.03664 0.0197 0.0824 1.0000
14.250 1.4031 0.04313 0.03702 0.0203 0.0798 1.0000
14.500 1.3993 0.04553 0.03960 0.0213 0.0777 1.0000
14.750 1.3898 0.04830 0.04265 0.0218 0.0757 1.0000
15.000 1.3854 0.05076 0.04528 0.0220 0.0735 1.0000
15.250 1.3915 0.05212 0.04664 0.0220 0.0711 1.0000
15.500 1.4018 0.05344 0.04782 0.0224 0.0684 1.0000
15.750 1.3848 0.05771 0.05242 0.0215 0.0672 1.0000
16.000 1.3667 0.06253 0.05751 0.0201 0.0659 1.0000
16.250 1.3492 0.06761 0.06281 0.0182 0.0645 1.0000
16.500 1.3344 0.07260 0.06796 0.0160 0.0631 1.0000
16.750 1.3421 0.07433 0.06963 0.0153 0.0608 1.0000
17.000 1.3399 0.07763 0.07289 0.0144 0.0588 1.0000
17.250 1.3089 0.08616 0.08169 0.0097 0.0584 1.0000
17.500 1.2693 0.09703 0.09283 0.0033 0.0584 1.0000
17.750 1.2107 0.11263 0.10867 -0.0061 0.0590 1.0000
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