DAE-21 AIRFOIL (dae21-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: DAE-21 AIRFOIL (dae21-il) Reynolds number: 500,000 Max Cl/Cd: 130.38 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-dae21-il-500000-n5.txt Download as CSV file: xf-dae21-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: DAE-21 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.0593 0.08943 0.08585 -0.0749 0.7204 0.0147
-8.500 -0.0545 0.08629 0.08271 -0.0765 0.7180 0.0147
-8.250 -0.0480 0.08343 0.07985 -0.0778 0.7158 0.0147
-8.000 -0.0414 0.08060 0.07700 -0.0790 0.7137 0.0147
-7.500 -0.0230 0.07595 0.07240 -0.0804 0.7091 0.0145
-7.250 -0.0148 0.07375 0.07021 -0.0811 0.7065 0.0137
-7.000 -0.0112 0.07099 0.06747 -0.0821 0.7037 0.0134
-6.750 -0.0100 0.06838 0.06487 -0.0829 0.7011 0.0133
-6.500 -0.0029 0.06549 0.06198 -0.0853 0.6988 0.0135
-6.250 0.0062 0.06214 0.05862 -0.0888 0.6964 0.0140
-6.000 0.0183 0.05848 0.05497 -0.0929 0.6938 0.0139
-5.750 0.0331 0.05444 0.05092 -0.0978 0.6911 0.0134
-5.500 0.0512 0.04873 0.04515 -0.1050 0.6884 0.0130
-5.250 0.0751 0.04056 0.03681 -0.1148 0.6858 0.0127
-5.000 0.1031 0.03070 0.02653 -0.1236 0.6836 0.0129
-4.750 0.1285 0.01764 0.01200 -0.1294 0.6817 0.0139
-4.500 0.1563 0.01574 0.00960 -0.1299 0.6794 0.0146
-4.250 0.1844 0.01469 0.00834 -0.1302 0.6768 0.0150
-4.000 0.2126 0.01406 0.00761 -0.1304 0.6740 0.0153
-3.750 0.2410 0.01352 0.00697 -0.1306 0.6713 0.0157
-3.500 0.2694 0.01300 0.00633 -0.1307 0.6686 0.0162
-3.250 0.2978 0.01252 0.00572 -0.1308 0.6661 0.0167
-3.000 0.3263 0.01208 0.00517 -0.1309 0.6636 0.0172
-2.750 0.3549 0.01171 0.00472 -0.1310 0.6607 0.0178
-2.500 0.3834 0.01133 0.00433 -0.1312 0.6577 0.0184
-2.250 0.4120 0.01111 0.00411 -0.1313 0.6546 0.0193
-2.000 0.4406 0.01090 0.00384 -0.1314 0.6517 0.0204
-1.750 0.4692 0.01068 0.00354 -0.1315 0.6490 0.0213
-1.500 0.4979 0.01043 0.00332 -0.1317 0.6460 0.0222
-1.250 0.5266 0.01026 0.00314 -0.1318 0.6426 0.0235
-1.000 0.5553 0.01012 0.00297 -0.1319 0.6392 0.0249
-0.750 0.5839 0.00998 0.00282 -0.1320 0.6359 0.0274
-0.250 0.6412 0.00976 0.00259 -0.1323 0.6293 0.0347
0.000 0.6698 0.00966 0.00252 -0.1324 0.6255 0.0439
0.250 0.6983 0.00954 0.00246 -0.1325 0.6216 0.0666
0.500 0.7265 0.00941 0.00243 -0.1327 0.6182 0.1110
0.750 0.7548 0.00921 0.00245 -0.1329 0.6143 0.1939
1.000 0.7829 0.00896 0.00251 -0.1332 0.6101 0.3134
1.250 0.8104 0.00870 0.00260 -0.1333 0.6058 0.4579
1.500 0.8369 0.00839 0.00269 -0.1332 0.6018 0.6237
1.750 0.8654 0.00771 0.00273 -0.1331 0.5973 1.0000
2.000 0.8936 0.00779 0.00276 -0.1332 0.5923 1.0000
2.250 0.9216 0.00789 0.00279 -0.1332 0.5874 1.0000
2.500 0.9496 0.00798 0.00284 -0.1333 0.5823 1.0000
2.750 0.9776 0.00807 0.00290 -0.1333 0.5765 1.0000
3.000 1.0051 0.00819 0.00296 -0.1333 0.5708 1.0000
3.250 1.0329 0.00829 0.00304 -0.1333 0.5647 1.0000
3.500 1.0603 0.00842 0.00313 -0.1333 0.5580 1.0000
3.750 1.0875 0.00855 0.00323 -0.1333 0.5513 1.0000
4.000 1.1146 0.00868 0.00334 -0.1332 0.5436 1.0000
4.250 1.1415 0.00884 0.00345 -0.1331 0.5361 1.0000
4.500 1.1681 0.00900 0.00359 -0.1330 0.5277 1.0000
4.750 1.1945 0.00917 0.00374 -0.1329 0.5195 1.0000
5.000 1.2205 0.00937 0.00390 -0.1327 0.5102 1.0000
5.250 1.2464 0.00956 0.00407 -0.1325 0.5010 1.0000
5.500 1.2715 0.00980 0.00428 -0.1321 0.4911 1.0000
5.750 1.2964 0.01004 0.00449 -0.1318 0.4799 1.0000
6.000 1.3208 0.01031 0.00472 -0.1314 0.4684 1.0000
6.250 1.3444 0.01061 0.00498 -0.1309 0.4562 1.0000
6.500 1.3671 0.01095 0.00528 -0.1302 0.4433 1.0000
7.000 1.4101 0.01171 0.00595 -0.1286 0.4152 1.0000
7.250 1.4299 0.01215 0.00634 -0.1275 0.3999 1.0000
7.500 1.4478 0.01265 0.00678 -0.1262 0.3836 1.0000
7.750 1.4640 0.01319 0.00726 -0.1246 0.3675 1.0000
8.000 1.4775 0.01379 0.00780 -0.1226 0.3510 1.0000
8.250 1.4857 0.01444 0.00841 -0.1196 0.3355 1.0000
8.500 1.4920 0.01524 0.00916 -0.1167 0.3201 1.0000
8.750 1.4996 0.01613 0.01000 -0.1142 0.3056 1.0000
9.000 1.5071 0.01713 0.01095 -0.1120 0.2907 1.0000
9.250 1.5146 0.01821 0.01199 -0.1098 0.2759 1.0000
9.500 1.5214 0.01939 0.01313 -0.1077 0.2608 1.0000
9.750 1.5280 0.02065 0.01433 -0.1058 0.2458 1.0000
10.000 1.5342 0.02198 0.01561 -0.1038 0.2309 1.0000
10.250 1.5401 0.02338 0.01696 -0.1019 0.2164 1.0000
10.500 1.5462 0.02480 0.01834 -0.1002 0.2025 1.0000
10.750 1.5531 0.02620 0.01971 -0.0985 0.1906 1.0000
11.000 1.5585 0.02774 0.02122 -0.0969 0.1778 1.0000
11.250 1.5641 0.02931 0.02275 -0.0953 0.1660 1.0000
11.500 1.5686 0.03100 0.02441 -0.0937 0.1537 1.0000
11.750 1.5723 0.03279 0.02616 -0.0922 0.1418 1.0000
12.000 1.5771 0.03452 0.02788 -0.0908 0.1314 1.0000
12.250 1.5821 0.03632 0.02966 -0.0895 0.1214 1.0000
12.500 1.5872 0.03816 0.03149 -0.0883 0.1124 1.0000
12.750 1.5911 0.04018 0.03349 -0.0871 0.1036 1.0000
13.000 1.5954 0.04219 0.03550 -0.0861 0.0953 1.0000
13.250 1.6008 0.04416 0.03749 -0.0852 0.0883 1.0000
13.500 1.6042 0.04638 0.03971 -0.0842 0.0817 1.0000
13.750 1.6091 0.04850 0.04184 -0.0835 0.0752 1.0000
14.000 1.6129 0.05077 0.04414 -0.0827 0.0700 1.0000
14.250 1.6169 0.05307 0.04647 -0.0821 0.0648 1.0000
14.500 1.6208 0.05544 0.04887 -0.0815 0.0605 1.0000
14.750 1.6238 0.05797 0.05142 -0.0810 0.0559 1.0000
15.000 1.6276 0.06044 0.05394 -0.0806 0.0525 1.0000
15.250 1.6303 0.06308 0.05663 -0.0803 0.0491 1.0000
15.500 1.6328 0.06579 0.05939 -0.0800 0.0462 1.0000
15.750 1.6356 0.06853 0.06220 -0.0799 0.0433 1.0000
16.000 1.6365 0.07156 0.06527 -0.0798 0.0407 1.0000
16.250 1.6389 0.07444 0.06822 -0.0798 0.0384 1.0000
16.500 1.6397 0.07756 0.07141 -0.0799 0.0361 1.0000
16.750 1.6394 0.08090 0.07481 -0.0800 0.0342 1.0000
17.000 1.6407 0.08407 0.07807 -0.0803 0.0324 1.0000
17.250 1.6407 0.08747 0.08154 -0.0807 0.0309 1.0000
17.500 1.6390 0.09114 0.08527 -0.0812 0.0294 1.0000
17.750 1.6385 0.09471 0.08893 -0.0818 0.0282 1.0000
18.000 1.6378 0.09835 0.09267 -0.0825 0.0270 1.0000
18.250 1.6361 0.10216 0.09657 -0.0833 0.0258 1.0000
18.500 1.6321 0.10640 0.10088 -0.0843 0.0248 1.0000
18.750 1.6307 0.11024 0.10482 -0.0853 0.0239 1.0000
19.000 1.6290 0.11417 0.10885 -0.0865 0.0229 1.0000
19.250 1.6260 0.11834 0.11311 -0.0878 0.0220 1.0000
|
Polar data table (+)
Polar graphs
<< Back to DAE-21 AIRFOIL (dae21-il)