NASA/LANGLEY LS(1)-0013 AIRFOIL (ls013-il) Xfoil prediction polar at RE=100,000 Ncrit=9
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Airfoil: NASA/LANGLEY LS(1)-0013 AIRFOIL (ls013-il) Reynolds number: 100,000 Max Cl/Cd: 36.59 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ls013-il-100000.txt Download as CSV file: xf-ls013-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY LS(1)-0013 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.8521 0.07900 0.07378 -0.0195 1.0000 0.0949
-10.750 -0.8682 0.07344 0.06813 -0.0204 1.0000 0.0931
-10.500 -0.9647 0.06356 0.05731 -0.0175 1.0000 0.0844
-10.250 -0.9670 0.06014 0.05346 -0.0160 1.0000 0.0836
-10.000 -0.9668 0.05551 0.04849 -0.0149 1.0000 0.0839
-9.750 -0.9592 0.05076 0.04351 -0.0143 1.0000 0.0849
-9.500 -0.9444 0.04737 0.03997 -0.0139 1.0000 0.0861
-9.250 -0.9306 0.04407 0.03636 -0.0132 1.0000 0.0869
-9.000 -0.9133 0.04136 0.03341 -0.0127 1.0000 0.0886
-8.750 -0.8949 0.03904 0.03080 -0.0121 1.0000 0.0914
-8.500 -0.8760 0.03664 0.02798 -0.0114 1.0000 0.0941
-8.250 -0.8553 0.03453 0.02536 -0.0107 1.0000 0.0963
-8.000 -0.8311 0.03196 0.02277 -0.0109 1.0000 0.1000
-7.750 -0.8065 0.03049 0.02122 -0.0109 1.0000 0.1047
-7.500 -0.7820 0.02901 0.01946 -0.0106 1.0000 0.1092
-7.250 -0.7571 0.02726 0.01772 -0.0105 1.0000 0.1144
-7.000 -0.7343 0.02615 0.01658 -0.0100 1.0000 0.1206
-6.750 -0.7140 0.02501 0.01532 -0.0089 1.0000 0.1265
-6.500 -0.6977 0.02402 0.01448 -0.0074 1.0000 0.1335
-6.250 -0.6823 0.02323 0.01358 -0.0054 1.0000 0.1410
-6.000 -0.6675 0.02231 0.01284 -0.0036 1.0000 0.1501
-5.750 -0.6531 0.02146 0.01209 -0.0017 1.0000 0.1615
-5.500 -0.6387 0.02062 0.01137 0.0002 1.0000 0.1764
-5.250 -0.6243 0.01974 0.01067 0.0020 1.0000 0.2002
-5.000 -0.6110 0.01856 0.00996 0.0038 1.0000 0.2471
-4.750 -0.6013 0.01661 0.00917 0.0056 1.0000 0.4010
-4.500 -0.5906 0.01615 0.00987 0.0097 1.0000 0.5987
-4.250 -0.5739 0.01666 0.01042 0.0125 1.0000 0.6677
-4.000 -0.5573 0.01733 0.01107 0.0155 1.0000 0.7048
-3.750 -0.5408 0.01797 0.01164 0.0184 1.0000 0.7337
-3.500 -0.5257 0.01869 0.01234 0.0220 1.0000 0.7555
-3.250 -0.5098 0.01931 0.01290 0.0253 0.9997 0.7765
-3.000 -0.4809 0.01997 0.01345 0.0260 0.9950 0.7987
-2.750 -0.4551 0.02063 0.01406 0.0278 0.9906 0.8170
-2.500 -0.4296 0.02122 0.01457 0.0295 0.9863 0.8353
-2.250 -0.3994 0.02187 0.01516 0.0307 0.9826 0.8555
-2.000 -0.3730 0.02229 0.01552 0.0321 0.9782 0.8756
-1.750 -0.3387 0.02259 0.01574 0.0316 0.9745 0.8916
-1.500 -0.2950 0.02278 0.01583 0.0285 0.9714 0.9021
-1.250 -0.2578 0.02284 0.01581 0.0262 0.9675 0.9128
-1.000 -0.2145 0.02295 0.01586 0.0227 0.9637 0.9210
-0.750 -0.1711 0.02300 0.01586 0.0190 0.9597 0.9289
-0.500 -0.1096 0.02321 0.01602 0.0120 0.9574 0.9367
-0.250 -0.0444 0.02335 0.01613 0.0042 0.9555 0.9439
0.000 0.0000 0.02339 0.01616 0.0000 0.9502 0.9502
0.250 0.0443 0.02335 0.01613 -0.0042 0.9439 0.9556
0.500 0.1096 0.02321 0.01602 -0.0120 0.9367 0.9574
0.750 0.1710 0.02300 0.01585 -0.0190 0.9289 0.9597
1.000 0.2145 0.02295 0.01585 -0.0227 0.9210 0.9637
1.250 0.2577 0.02283 0.01580 -0.0262 0.9128 0.9675
1.500 0.2950 0.02278 0.01583 -0.0285 0.9021 0.9714
1.750 0.3387 0.02259 0.01573 -0.0316 0.8917 0.9746
2.000 0.3730 0.02229 0.01552 -0.0321 0.8756 0.9782
2.250 0.3994 0.02187 0.01516 -0.0307 0.8555 0.9826
2.500 0.4295 0.02121 0.01457 -0.0295 0.8353 0.9863
2.750 0.4550 0.02063 0.01405 -0.0277 0.8170 0.9906
3.000 0.4808 0.01996 0.01345 -0.0260 0.7987 0.9950
3.250 0.5098 0.01931 0.01289 -0.0253 0.7766 0.9997
3.500 0.5256 0.01869 0.01234 -0.0220 0.7556 1.0000
3.750 0.5406 0.01797 0.01164 -0.0184 0.7338 1.0000
4.000 0.5571 0.01733 0.01107 -0.0155 0.7049 1.0000
4.250 0.5737 0.01666 0.01042 -0.0124 0.6677 1.0000
4.500 0.5905 0.01614 0.00987 -0.0096 0.5990 1.0000
4.750 0.6012 0.01659 0.00917 -0.0056 0.4018 1.0000
5.000 0.6109 0.01855 0.00995 -0.0037 0.2473 1.0000
5.250 0.6241 0.01974 0.01067 -0.0020 0.2002 1.0000
5.500 0.6385 0.02061 0.01137 -0.0002 0.1765 1.0000
5.750 0.6529 0.02146 0.01208 0.0017 0.1615 1.0000
6.000 0.6674 0.02231 0.01283 0.0036 0.1501 1.0000
6.250 0.6821 0.02322 0.01357 0.0054 0.1411 1.0000
6.500 0.6975 0.02402 0.01448 0.0074 0.1335 1.0000
6.750 0.7140 0.02501 0.01532 0.0090 0.1265 1.0000
7.000 0.7343 0.02615 0.01657 0.0100 0.1206 1.0000
7.250 0.7571 0.02726 0.01772 0.0105 0.1144 1.0000
7.500 0.7820 0.02901 0.01946 0.0106 0.1092 1.0000
7.750 0.8065 0.03049 0.02122 0.0109 0.1047 1.0000
8.000 0.8312 0.03196 0.02277 0.0109 0.1000 1.0000
8.250 0.8553 0.03453 0.02536 0.0107 0.0963 1.0000
8.500 0.8761 0.03664 0.02798 0.0114 0.0941 1.0000
8.750 0.8950 0.03904 0.03080 0.0121 0.0914 1.0000
9.000 0.9134 0.04136 0.03341 0.0126 0.0885 1.0000
9.250 0.9308 0.04408 0.03636 0.0132 0.0869 1.0000
9.500 0.9445 0.04739 0.04000 0.0139 0.0861 1.0000
9.750 0.9593 0.05077 0.04353 0.0143 0.0849 1.0000
10.000 0.9670 0.05549 0.04847 0.0148 0.0839 1.0000
10.250 0.9675 0.06020 0.05351 0.0159 0.0837 1.0000
10.500 0.9646 0.06354 0.05730 0.0175 0.0845 1.0000
10.750 0.8687 0.07349 0.06818 0.0203 0.0932 1.0000
11.000 0.8530 0.07904 0.07382 0.0194 0.0949 1.0000
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Polar data table (+)
Polar graphs
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