Davis AIRFOIL (davis-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: Davis AIRFOIL (davis-il) Reynolds number: 50,000 Max Cl/Cd: 36.44 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-davis-il-50000-n5.txt Download as CSV file: xf-davis-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: Davis AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.3705 0.10228 0.09545 -0.0435 1.0000 0.0530
-9.250 -0.3774 0.09856 0.09183 -0.0440 1.0000 0.0521
-9.000 -0.3871 0.09468 0.08807 -0.0445 1.0000 0.0513
-8.750 -0.4017 0.09051 0.08405 -0.0451 1.0000 0.0502
-8.500 -0.4355 0.08426 0.07797 -0.0478 1.0000 0.0482
-8.250 -0.4504 0.08093 0.07476 -0.0472 1.0000 0.0480
-8.000 -0.4627 0.07704 0.07093 -0.0473 1.0000 0.0477
-7.750 -0.4755 0.07281 0.06674 -0.0475 1.0000 0.0475
-7.500 -0.4879 0.06875 0.06268 -0.0473 1.0000 0.0472
-7.250 -0.4994 0.06441 0.05826 -0.0472 1.0000 0.0471
-7.000 -0.5080 0.05991 0.05359 -0.0471 1.0000 0.0469
-6.750 -0.5127 0.05532 0.04874 -0.0470 1.0000 0.0469
-6.500 -0.5132 0.05063 0.04363 -0.0468 1.0000 0.0472
-6.250 -0.5086 0.04620 0.03857 -0.0465 1.0000 0.0481
-6.000 -0.4988 0.04383 0.03612 -0.0455 1.0000 0.0501
-5.750 -0.4795 0.04147 0.03347 -0.0460 0.9976 0.0532
-5.500 -0.4506 0.03800 0.02935 -0.0480 0.9924 0.0562
-5.250 -0.4184 0.03502 0.02555 -0.0497 0.9871 0.0601
-5.000 -0.3867 0.03343 0.02387 -0.0515 0.9803 0.0665
-4.750 -0.3522 0.03150 0.02134 -0.0530 0.9736 0.0731
-4.500 -0.3176 0.03036 0.02011 -0.0549 0.9667 0.0843
-4.250 -0.2842 0.02915 0.01875 -0.0562 0.9595 0.0956
-4.000 -0.2521 0.02825 0.01771 -0.0573 0.9522 0.1162
-3.750 -0.2157 0.02733 0.01681 -0.0592 0.9466 0.1513
-3.500 -0.1868 0.02678 0.01619 -0.0598 0.9388 0.1914
-3.250 -0.1560 0.02670 0.01612 -0.0610 0.9316 0.2427
-3.000 -0.1259 0.02652 0.01596 -0.0619 0.9243 0.2884
-2.750 -0.0959 0.02622 0.01557 -0.0625 0.9169 0.3191
-2.500 -0.0626 0.02592 0.01521 -0.0637 0.9102 0.3521
-2.250 -0.0340 0.02567 0.01495 -0.0641 0.9025 0.3865
-2.000 -0.0012 0.02535 0.01465 -0.0652 0.8957 0.4284
-1.750 0.0261 0.02505 0.01444 -0.0652 0.8877 0.4777
-1.500 0.0566 0.02456 0.01424 -0.0657 0.8810 0.5475
-1.250 0.0818 0.02390 0.01424 -0.0644 0.8737 0.6843
-1.000 0.1558 0.02361 0.01406 -0.0731 0.8699 1.0000
-0.750 0.1944 0.02369 0.01380 -0.0753 0.8639 1.0000
-0.500 0.2166 0.02391 0.01379 -0.0746 0.8531 1.0000
-0.250 0.2463 0.02407 0.01373 -0.0751 0.8447 1.0000
0.000 0.2779 0.02418 0.01365 -0.0760 0.8366 1.0000
0.250 0.3038 0.02435 0.01366 -0.0757 0.8262 1.0000
0.500 0.3444 0.02424 0.01338 -0.0778 0.8194 1.0000
0.750 0.3710 0.02428 0.01331 -0.0775 0.8071 1.0000
1.000 0.4016 0.02420 0.01311 -0.0776 0.7949 1.0000
1.250 0.4342 0.02402 0.01284 -0.0779 0.7829 1.0000
1.500 0.4701 0.02378 0.01250 -0.0788 0.7725 1.0000
1.750 0.5016 0.02369 0.01237 -0.0790 0.7617 1.0000
2.000 0.5274 0.02381 0.01244 -0.0784 0.7499 1.0000
2.250 0.5559 0.02386 0.01246 -0.0783 0.7391 1.0000
2.500 0.5926 0.02368 0.01225 -0.0793 0.7303 1.0000
2.750 0.6170 0.02384 0.01242 -0.0785 0.7174 1.0000
3.000 0.6420 0.02399 0.01257 -0.0777 0.7044 1.0000
3.250 0.6675 0.02411 0.01273 -0.0770 0.6912 1.0000
3.500 0.6930 0.02424 0.01287 -0.0763 0.6776 1.0000
3.750 0.7179 0.02436 0.01302 -0.0754 0.6635 1.0000
4.000 0.7420 0.02451 0.01323 -0.0745 0.6488 1.0000
4.250 0.7655 0.02466 0.01343 -0.0734 0.6335 1.0000
4.500 0.7889 0.02481 0.01363 -0.0723 0.6178 1.0000
4.750 0.8131 0.02495 0.01385 -0.0713 0.6020 1.0000
5.000 0.8382 0.02509 0.01404 -0.0704 0.5863 1.0000
5.250 0.8640 0.02524 0.01423 -0.0697 0.5706 1.0000
5.500 0.8901 0.02545 0.01448 -0.0690 0.5549 1.0000
5.750 0.9158 0.02574 0.01485 -0.0684 0.5394 1.0000
6.000 0.9408 0.02613 0.01529 -0.0677 0.5242 1.0000
6.250 0.9649 0.02661 0.01586 -0.0670 0.5096 1.0000
6.500 0.9881 0.02718 0.01654 -0.0662 0.4957 1.0000
6.750 1.0106 0.02780 0.01733 -0.0653 0.4825 1.0000
7.000 1.0334 0.02844 0.01812 -0.0645 0.4703 1.0000
7.250 1.0583 0.02904 0.01886 -0.0640 0.4594 1.0000
7.500 1.0775 0.02985 0.01989 -0.0628 0.4485 1.0000
7.750 1.0970 0.03067 0.02094 -0.0616 0.4383 1.0000
8.000 1.1214 0.03134 0.02186 -0.0610 0.4293 1.0000
8.250 1.1388 0.03218 0.02298 -0.0595 0.4190 1.0000
8.500 1.1537 0.03248 0.02339 -0.0571 0.4023 1.0000
8.750 1.1563 0.03287 0.02384 -0.0529 0.3800 1.0000
9.000 1.1563 0.03316 0.02403 -0.0482 0.3539 1.0000
9.250 1.1495 0.03384 0.02454 -0.0429 0.3263 1.0000
9.500 1.1387 0.03503 0.02566 -0.0380 0.2971 1.0000
9.750 1.1297 0.03654 0.02721 -0.0340 0.2667 1.0000
10.000 1.1232 0.03833 0.02912 -0.0309 0.2306 1.0000
10.250 1.1134 0.04068 0.03129 -0.0281 0.1660 1.0000
10.500 1.0957 0.04420 0.03418 -0.0256 0.1131 1.0000
10.750 1.0801 0.04821 0.03789 -0.0240 0.0891 1.0000
11.000 1.0677 0.05221 0.04184 -0.0229 0.0760 1.0000
11.250 1.0583 0.05609 0.04577 -0.0222 0.0681 1.0000
11.500 1.0481 0.06022 0.04995 -0.0219 0.0622 1.0000
11.750 1.0401 0.06423 0.05403 -0.0218 0.0581 1.0000
12.000 1.0343 0.06809 0.05810 -0.0219 0.0537 1.0000
12.250 1.0289 0.07200 0.06210 -0.0222 0.0508 1.0000
12.500 1.0220 0.07615 0.06624 -0.0226 0.0483 1.0000
12.750 1.0224 0.07941 0.06971 -0.0225 0.0456 1.0000
13.000 1.0243 0.08245 0.07290 -0.0223 0.0437 1.0000
13.250 1.0269 0.08543 0.07600 -0.0220 0.0417 1.0000
13.500 1.0298 0.08838 0.07901 -0.0219 0.0398 1.0000
13.750 1.0345 0.09108 0.08174 -0.0215 0.0379 1.0000
14.000 1.0407 0.09394 0.08489 -0.0212 0.0361 1.0000
14.250 1.0453 0.09714 0.08832 -0.0213 0.0345 1.0000
14.500 1.0499 0.10042 0.09184 -0.0213 0.0337 1.0000
14.750 1.0518 0.10427 0.09591 -0.0219 0.0330 1.0000
15.000 1.0506 0.10870 0.10058 -0.0230 0.0326 1.0000
15.250 1.0461 0.11382 0.10592 -0.0249 0.0323 1.0000
15.500 1.0383 0.11968 0.11201 -0.0275 0.0321 1.0000
15.750 1.0275 0.12637 0.11891 -0.0309 0.0321 1.0000
16.000 1.0133 0.13422 0.12699 -0.0353 0.0323 1.0000
16.250 0.9932 0.14424 0.13723 -0.0414 0.0329 1.0000
16.500 0.9714 0.15578 0.14890 -0.0484 0.0337 1.0000
|
Polar data table (+)
Polar graphs
<< Back to Davis AIRFOIL (davis-il)