Falcon (falcon-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: Falcon (falcon-il) Reynolds number: 50,000 Max Cl/Cd: 30.2 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-falcon-il-50000-n5.txt Download as CSV file: xf-falcon-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: Falcon
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
------ -------- --------- --------- -------- -------- --------
-11.000 -0.5709 0.09761 0.09012 -0.0285 1.0000 0.0635
-10.750 -0.5889 0.08926 0.08179 -0.0338 1.0000 0.0630
-10.500 -0.6145 0.08148 0.07400 -0.0388 1.0000 0.0624
-10.250 -0.6426 0.07523 0.06768 -0.0414 1.0000 0.0619
-10.000 -0.6690 0.07048 0.06282 -0.0411 1.0000 0.0616
-9.750 -0.6924 0.06653 0.05874 -0.0389 1.0000 0.0616
-9.500 -0.7096 0.06261 0.05460 -0.0366 1.0000 0.0617
-9.250 -0.7220 0.05886 0.05058 -0.0340 1.0000 0.0622
-9.000 -0.7308 0.05527 0.04664 -0.0311 1.0000 0.0631
-8.750 -0.7368 0.05177 0.04268 -0.0279 1.0000 0.0645
-8.500 -0.7340 0.04895 0.03957 -0.0253 1.0000 0.0661
-8.250 -0.7230 0.04706 0.03761 -0.0235 1.0000 0.0681
-8.000 -0.7128 0.04485 0.03520 -0.0214 1.0000 0.0701
-7.750 -0.7020 0.04247 0.03252 -0.0192 1.0000 0.0721
-7.500 -0.6903 0.04021 0.02982 -0.0169 1.0000 0.0754
-7.250 -0.6755 0.03832 0.02772 -0.0150 1.0000 0.0788
-7.000 -0.6586 0.03687 0.02622 -0.0135 1.0000 0.0824
-6.750 -0.6406 0.03522 0.02431 -0.0119 1.0000 0.0868
-6.500 -0.6222 0.03373 0.02266 -0.0103 1.0000 0.0921
-6.250 -0.6034 0.03256 0.02146 -0.0089 1.0000 0.0978
-6.000 -0.5829 0.03129 0.01996 -0.0075 1.0000 0.1049
-5.750 -0.5643 0.03028 0.01903 -0.0061 1.0000 0.1126
-5.500 -0.5442 0.02924 0.01790 -0.0047 1.0000 0.1217
-5.250 -0.5256 0.02838 0.01698 -0.0031 1.0000 0.1327
-5.000 -0.5078 0.02750 0.01616 -0.0015 1.0000 0.1443
-4.750 -0.4906 0.02668 0.01536 0.0002 1.0000 0.1577
-4.500 -0.4743 0.02591 0.01461 0.0019 1.0000 0.1733
-4.000 -0.4439 0.02439 0.01324 0.0054 1.0000 0.2148
-3.750 -0.4300 0.02350 0.01270 0.0071 1.0000 0.2484
-3.500 -0.4042 0.02203 0.01222 0.0063 0.9910 0.3835
-3.250 -0.3723 0.02116 0.01221 0.0059 0.9777 0.5698
-3.000 -0.3320 0.02106 0.01244 0.0048 0.9658 0.6805
-2.750 -0.2860 0.02127 0.01271 0.0029 0.9548 0.7587
-2.500 -0.2372 0.02157 0.01291 0.0002 0.9431 0.8148
-2.250 -0.1788 0.02197 0.01315 -0.0041 0.9326 0.8562
-2.000 -0.1111 0.02243 0.01338 -0.0103 0.9235 0.8940
-1.750 -0.0378 0.02263 0.01334 -0.0182 0.9123 0.9164
-1.500 0.0189 0.02249 0.01300 -0.0238 0.8954 0.9295
-1.250 0.0694 0.02233 0.01266 -0.0284 0.8748 0.9417
-1.000 0.1216 0.02213 0.01229 -0.0332 0.8534 0.9522
-0.750 0.1712 0.02191 0.01189 -0.0375 0.8319 0.9631
-0.500 0.2168 0.02172 0.01155 -0.0411 0.8084 0.9744
-0.250 0.2611 0.02155 0.01122 -0.0444 0.7856 0.9860
0.000 0.3049 0.02139 0.01089 -0.0478 0.7634 0.9973
0.250 0.3314 0.02130 0.01066 -0.0479 0.7414 1.0000
0.500 0.3519 0.02125 0.01048 -0.0468 0.7215 1.0000
0.750 0.3725 0.02124 0.01033 -0.0456 0.7026 1.0000
1.000 0.3932 0.02126 0.01022 -0.0444 0.6848 1.0000
1.250 0.4137 0.02132 0.01017 -0.0432 0.6675 1.0000
1.500 0.4340 0.02142 0.01017 -0.0419 0.6507 1.0000
1.750 0.4543 0.02154 0.01020 -0.0406 0.6346 1.0000
2.000 0.4747 0.02170 0.01028 -0.0394 0.6191 1.0000
2.250 0.4954 0.02188 0.01040 -0.0382 0.6036 1.0000
2.500 0.5161 0.02210 0.01056 -0.0370 0.5884 1.0000
2.750 0.5367 0.02233 0.01075 -0.0358 0.5737 1.0000
3.000 0.5573 0.02259 0.01098 -0.0346 0.5592 1.0000
3.250 0.5778 0.02286 0.01122 -0.0333 0.5453 1.0000
3.500 0.5982 0.02315 0.01146 -0.0320 0.5316 1.0000
3.750 0.6191 0.02343 0.01171 -0.0307 0.5185 1.0000
4.000 0.6397 0.02373 0.01196 -0.0294 0.5052 1.0000
4.250 0.6594 0.02409 0.01233 -0.0280 0.4913 1.0000
4.500 0.6791 0.02446 0.01270 -0.0266 0.4777 1.0000
4.750 0.6992 0.02484 0.01309 -0.0252 0.4646 1.0000
5.000 0.7196 0.02521 0.01345 -0.0239 0.4520 1.0000
5.250 0.7406 0.02556 0.01375 -0.0226 0.4399 1.0000
5.500 0.7596 0.02602 0.01426 -0.0211 0.4266 1.0000
5.750 0.7789 0.02650 0.01481 -0.0197 0.4138 1.0000
6.000 0.7984 0.02698 0.01531 -0.0183 0.4015 1.0000
6.250 0.8187 0.02743 0.01575 -0.0169 0.3899 1.0000
6.500 0.8381 0.02795 0.01633 -0.0155 0.3781 1.0000
6.750 0.8562 0.02856 0.01702 -0.0140 0.3659 1.0000
7.000 0.8751 0.02916 0.01767 -0.0126 0.3547 1.0000
7.250 0.8958 0.02966 0.01813 -0.0114 0.3446 1.0000
7.500 0.9122 0.03044 0.01908 -0.0097 0.3330 1.0000
7.750 0.9300 0.03118 0.01991 -0.0083 0.3227 1.0000
8.000 0.9503 0.03179 0.02051 -0.0071 0.3137 1.0000
8.250 0.9653 0.03276 0.02165 -0.0054 0.3035 1.0000
8.500 0.9840 0.03351 0.02246 -0.0041 0.2950 1.0000
8.750 0.9997 0.03448 0.02358 -0.0025 0.2862 1.0000
9.000 1.0172 0.03543 0.02463 -0.0012 0.2788 1.0000
9.250 1.0321 0.03651 0.02586 0.0004 0.2710 1.0000
9.500 1.0495 0.03748 0.02691 0.0017 0.2643 1.0000
9.750 1.0609 0.03880 0.02848 0.0035 0.2570 1.0000
10.000 1.0830 0.03954 0.02921 0.0043 0.2514 1.0000
10.250 1.0868 0.04146 0.03148 0.0066 0.2452 1.0000
10.500 1.1003 0.04274 0.03290 0.0081 0.2397 1.0000
10.750 1.1160 0.04391 0.03420 0.0094 0.2347 1.0000
11.000 1.1128 0.04617 0.03678 0.0121 0.2293 1.0000
11.250 1.1236 0.04758 0.03834 0.0137 0.2243 1.0000
11.500 1.1347 0.04894 0.03982 0.0153 0.2195 1.0000
11.750 1.1132 0.05207 0.04323 0.0190 0.2153 1.0000
12.000 1.1374 0.05130 0.04237 0.0203 0.2056 1.0000
12.250 1.1063 0.05554 0.04691 0.0232 0.2025 1.0000
12.500 1.0717 0.06113 0.05272 0.0240 0.1998 1.0000
13.000 1.0420 0.06890 0.06066 0.0236 0.1881 1.0000
13.250 0.9682 0.08430 0.07613 0.0174 0.1909 1.0000
13.500 0.8205 0.11595 0.10738 0.0007 0.1841 1.0000
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Polar data table (+)
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