USA 40 AIRFOIL (usa40-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: USA 40 AIRFOIL (usa40-il) Reynolds number: 500,000 Max Cl/Cd: 82.19 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa40-il-500000-n5.txt Download as CSV file: xf-usa40-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA 40 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.000 -0.7565 0.04095 0.03753 -0.0834 1.0000 0.0254
-11.750 -0.7645 0.03533 0.03162 -0.0913 0.9962 0.0255
-11.500 -0.7434 0.03165 0.02762 -0.0973 0.9909 0.0257
-11.250 -0.7193 0.02906 0.02478 -0.1009 0.9861 0.0259
-11.000 -0.6913 0.02698 0.02246 -0.1041 0.9830 0.0261
-10.750 -0.6654 0.02535 0.02062 -0.1059 0.9779 0.0263
-10.500 -0.6367 0.02393 0.01902 -0.1078 0.9735 0.0265
-10.250 -0.6060 0.02270 0.01761 -0.1097 0.9702 0.0266
-10.000 -0.5807 0.02142 0.01620 -0.1104 0.9634 0.0268
-9.750 -0.5512 0.02025 0.01493 -0.1118 0.9586 0.0271
-9.500 -0.5224 0.01935 0.01396 -0.1127 0.9526 0.0273
-9.250 -0.4932 0.01857 0.01311 -0.1136 0.9458 0.0276
-9.000 -0.4628 0.01785 0.01232 -0.1145 0.9399 0.0278
-8.750 -0.4345 0.01724 0.01164 -0.1150 0.9319 0.0282
-8.250 -0.3767 0.01609 0.01034 -0.1158 0.9149 0.0289
-8.000 -0.3481 0.01553 0.00968 -0.1161 0.9055 0.0293
-7.750 -0.3206 0.01501 0.00906 -0.1162 0.8929 0.0296
-7.500 -0.2933 0.01453 0.00847 -0.1161 0.8772 0.0299
-7.250 -0.2662 0.01411 0.00793 -0.1159 0.8603 0.0302
-7.000 -0.2390 0.01372 0.00742 -0.1157 0.8449 0.0305
-6.750 -0.2116 0.01337 0.00696 -0.1156 0.8311 0.0308
-6.500 -0.1844 0.01297 0.00647 -0.1155 0.8165 0.0311
-6.250 -0.1571 0.01257 0.00601 -0.1154 0.8013 0.0316
-6.000 -0.1297 0.01226 0.00563 -0.1153 0.7858 0.0321
-5.750 -0.1021 0.01200 0.00529 -0.1151 0.7701 0.0326
-5.500 -0.0745 0.01177 0.00497 -0.1150 0.7537 0.0331
-5.250 -0.0469 0.01157 0.00468 -0.1148 0.7370 0.0338
-5.000 -0.0193 0.01139 0.00440 -0.1147 0.7203 0.0346
-4.750 0.0085 0.01124 0.00415 -0.1145 0.7040 0.0353
-4.500 0.0361 0.01106 0.00389 -0.1144 0.6874 0.0363
-4.250 0.0639 0.01090 0.00366 -0.1142 0.6719 0.0374
-4.000 0.0920 0.01077 0.00346 -0.1142 0.6581 0.0386
-3.750 0.1200 0.01066 0.00328 -0.1140 0.6447 0.0400
-3.500 0.1480 0.01056 0.00311 -0.1139 0.6315 0.0419
-3.250 0.1762 0.01045 0.00297 -0.1139 0.6186 0.0447
-3.000 0.2045 0.01036 0.00285 -0.1138 0.6074 0.0488
-2.750 0.2325 0.01029 0.00275 -0.1137 0.5950 0.0552
-2.500 0.2608 0.01022 0.00267 -0.1137 0.5816 0.0635
-2.250 0.2890 0.01016 0.00259 -0.1136 0.5687 0.0717
-2.000 0.3171 0.01013 0.00251 -0.1135 0.5556 0.0794
-1.750 0.3451 0.01009 0.00244 -0.1135 0.5405 0.0885
-1.500 0.3731 0.01005 0.00238 -0.1134 0.5235 0.1019
-1.250 0.4010 0.00994 0.00234 -0.1134 0.5062 0.1384
-1.000 0.4288 0.00990 0.00235 -0.1134 0.4897 0.1801
-0.750 0.4564 0.00992 0.00236 -0.1133 0.4713 0.2056
-0.500 0.4837 0.00997 0.00238 -0.1131 0.4517 0.2320
0.000 0.5389 0.00993 0.00247 -0.1131 0.4185 0.3247
0.250 0.5667 0.00990 0.00253 -0.1131 0.4063 0.3824
0.500 0.5944 0.00993 0.00260 -0.1130 0.3950 0.4224
0.750 0.6224 0.00988 0.00268 -0.1130 0.3850 0.4848
1.000 0.6500 0.00985 0.00280 -0.1130 0.3756 0.5531
1.250 0.6774 0.00992 0.00292 -0.1128 0.3642 0.5995
1.500 0.7047 0.01000 0.00306 -0.1125 0.3537 0.6417
1.750 0.7316 0.01014 0.00320 -0.1122 0.3424 0.6741
2.000 0.7589 0.01027 0.00331 -0.1120 0.3311 0.6931
2.250 0.7860 0.01043 0.00345 -0.1117 0.3209 0.7104
2.500 0.8129 0.01059 0.00358 -0.1114 0.3108 0.7279
2.750 0.8400 0.01074 0.00372 -0.1112 0.3022 0.7430
3.000 0.8668 0.01093 0.00386 -0.1109 0.2924 0.7543
3.250 0.8936 0.01110 0.00401 -0.1106 0.2809 0.7658
3.500 0.9198 0.01131 0.00417 -0.1102 0.2682 0.7769
4.000 0.9711 0.01184 0.00455 -0.1094 0.2357 0.7985
4.250 0.9961 0.01212 0.00477 -0.1089 0.2191 0.8128
4.500 1.0202 0.01246 0.00503 -0.1082 0.1993 0.8291
4.750 1.0428 0.01286 0.00534 -0.1073 0.1750 0.8502
5.250 1.0791 0.01439 0.00651 -0.1042 0.0849 1.0000
5.500 1.1039 0.01477 0.00685 -0.1037 0.0791 1.0000
5.750 1.1288 0.01513 0.00717 -0.1033 0.0741 1.0000
6.000 1.1533 0.01550 0.00751 -0.1028 0.0687 1.0000
6.250 1.1755 0.01606 0.00795 -0.1020 0.0413 1.0000
6.500 1.1971 0.01667 0.00849 -0.1011 0.0334 1.0000
6.750 1.2204 0.01708 0.00892 -0.1004 0.0322 1.0000
7.000 1.2432 0.01751 0.00936 -0.0997 0.0312 1.0000
7.250 1.2655 0.01797 0.00983 -0.0989 0.0305 1.0000
7.500 1.2871 0.01844 0.01033 -0.0980 0.0298 1.0000
7.750 1.3086 0.01889 0.01081 -0.0970 0.0293 1.0000
8.000 1.3294 0.01937 0.01132 -0.0960 0.0288 1.0000
8.250 1.3493 0.01988 0.01187 -0.0949 0.0283 1.0000
8.500 1.3672 0.02042 0.01245 -0.0934 0.0279 1.0000
8.750 1.3838 0.02101 0.01308 -0.0918 0.0275 1.0000
9.000 1.3992 0.02167 0.01378 -0.0901 0.0272 1.0000
9.250 1.4132 0.02243 0.01459 -0.0882 0.0269 1.0000
9.500 1.4248 0.02334 0.01556 -0.0862 0.0265 1.0000
9.750 1.4386 0.02413 0.01639 -0.0845 0.0262 1.0000
10.000 1.4515 0.02499 0.01731 -0.0828 0.0259 1.0000
10.250 1.4633 0.02596 0.01834 -0.0811 0.0256 1.0000
10.500 1.4737 0.02708 0.01951 -0.0795 0.0253 1.0000
10.750 1.4830 0.02832 0.02082 -0.0779 0.0250 1.0000
11.000 1.4910 0.02973 0.02229 -0.0763 0.0248 1.0000
11.250 1.4980 0.03128 0.02392 -0.0749 0.0245 1.0000
11.500 1.5036 0.03305 0.02576 -0.0736 0.0243 1.0000
11.750 1.5078 0.03504 0.02783 -0.0725 0.0241 1.0000
12.000 1.5101 0.03730 0.03018 -0.0715 0.0238 1.0000
12.250 1.5103 0.03992 0.03288 -0.0707 0.0236 1.0000
12.500 1.5076 0.04299 0.03604 -0.0702 0.0234 1.0000
12.750 1.5018 0.04656 0.03971 -0.0699 0.0232 1.0000
13.000 1.5012 0.04964 0.04290 -0.0698 0.0230 1.0000
13.250 1.4997 0.05293 0.04629 -0.0699 0.0229 1.0000
13.500 1.4966 0.05649 0.04995 -0.0701 0.0227 1.0000
13.750 1.4922 0.06030 0.05387 -0.0705 0.0226 1.0000
14.000 1.4864 0.06441 0.05809 -0.0710 0.0225 1.0000
14.250 1.4797 0.06870 0.06248 -0.0717 0.0223 1.0000
14.500 1.4726 0.07313 0.06702 -0.0725 0.0221 1.0000
14.750 1.4650 0.07774 0.07173 -0.0734 0.0220 1.0000
15.000 1.4573 0.08240 0.07649 -0.0744 0.0218 1.0000
15.250 1.4497 0.08707 0.08126 -0.0754 0.0216 1.0000
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