FX 63-143 AIRFOIL (fx63143-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: FX 63-143 AIRFOIL (fx63143-il) Reynolds number: 1,000,000 Max Cl/Cd: 79.43 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx63143-il-1000000.txt Download as CSV file: xf-fx63143-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: FX 63-143 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.2678 0.08649 0.08427 -0.0757 0.8337 0.0125
-10.500 -0.2703 0.08231 0.07971 -0.0777 0.7493 0.0127
-10.250 -0.2753 0.07703 0.07437 -0.0806 0.7291 0.0130
-10.000 -0.2823 0.07107 0.06841 -0.0845 0.7242 0.0129
-9.250 -0.3572 0.03584 0.03270 -0.0894 0.6779 0.0135
-9.000 -0.3777 0.03351 0.03023 -0.0854 0.6701 0.0135
-8.750 -0.4011 0.03059 0.02705 -0.0807 0.6650 0.0136
-8.500 -0.4138 0.02761 0.02377 -0.0768 0.6575 0.0136
-8.250 -0.4247 0.02467 0.02045 -0.0726 0.6503 0.0136
-7.500 -0.3978 0.02897 0.02429 -0.0721 0.6272 0.0139
-7.250 -0.3812 0.02768 0.02292 -0.0711 0.6196 0.0141
-7.000 -0.3652 0.02627 0.02136 -0.0697 0.6121 0.0145
-6.750 -0.3490 0.02451 0.01937 -0.0680 0.6056 0.0148
-6.500 -0.3348 0.02299 0.01724 -0.0650 0.5983 0.0158
-6.250 -0.3190 0.02207 0.01583 -0.0629 0.5919 0.0159
-6.000 -0.2942 0.01820 0.01181 -0.0622 0.5853 0.0139
-5.750 -0.2690 0.01561 0.00895 -0.0612 0.5755 0.0114
-5.500 -0.2444 0.01429 0.00762 -0.0607 0.5653 0.0110
-5.250 -0.2287 0.01204 0.00526 -0.0584 0.5571 0.0097
-5.000 -0.2076 0.01130 0.00442 -0.0572 0.5482 0.0093
-4.750 -0.1840 0.01088 0.00388 -0.0564 0.5390 0.0089
-4.500 -0.1585 0.01064 0.00354 -0.0560 0.5306 0.0086
-4.250 -0.1326 0.01048 0.00329 -0.0556 0.5199 0.0084
-4.000 -0.1065 0.01036 0.00310 -0.0553 0.5073 0.0082
-3.750 -0.0798 0.01025 0.00294 -0.0551 0.4966 0.0080
-3.500 -0.0530 0.01019 0.00284 -0.0550 0.4865 0.0079
-3.250 -0.0264 0.01015 0.00273 -0.0548 0.4750 0.0078
-3.000 0.0005 0.01009 0.00264 -0.0546 0.4628 0.0076
-2.750 0.0273 0.01005 0.00256 -0.0545 0.4495 0.0075
-2.500 0.0534 0.01007 0.00248 -0.0543 0.4274 0.0075
-2.250 0.0799 0.01007 0.00241 -0.0541 0.4130 0.0075
-2.000 0.1065 0.01006 0.00233 -0.0539 0.3987 0.0076
-1.750 0.1318 0.01016 0.00232 -0.0536 0.3643 0.0081
-1.500 0.1573 0.01029 0.00232 -0.0533 0.3399 0.0115
-1.250 0.1813 0.01054 0.00237 -0.0527 0.2917 0.0128
-1.000 0.2058 0.01078 0.00245 -0.0522 0.2474 0.0134
-0.750 0.2306 0.01100 0.00254 -0.0519 0.2194 0.0144
-0.500 0.2566 0.01108 0.00256 -0.0517 0.1915 0.0149
-0.250 0.2822 0.01122 0.00260 -0.0514 0.1726 0.0153
0.000 0.3089 0.01132 0.00267 -0.0513 0.1634 0.0159
0.250 0.3349 0.01135 0.00263 -0.0512 0.1389 0.0164
0.500 0.3612 0.01141 0.00263 -0.0510 0.1304 0.0168
0.750 0.3882 0.01143 0.00264 -0.0510 0.1253 0.0177
1.000 0.4149 0.01150 0.00268 -0.0510 0.1179 0.0186
1.250 0.4420 0.01151 0.00265 -0.0510 0.1127 0.0208
1.500 0.4681 0.01162 0.00271 -0.0508 0.0951 0.0308
1.750 0.4947 0.01173 0.00279 -0.0508 0.0914 0.0333
2.000 0.5217 0.01180 0.00286 -0.0508 0.0891 0.0357
2.250 0.5488 0.01187 0.00294 -0.0508 0.0875 0.0371
2.500 0.5758 0.01193 0.00301 -0.0508 0.0869 0.0377
2.750 0.6029 0.01200 0.00308 -0.0509 0.0866 0.0389
3.000 0.6299 0.01207 0.00317 -0.0509 0.0864 0.0404
3.250 0.6569 0.01215 0.00326 -0.0509 0.0861 0.0432
3.500 0.6838 0.01222 0.00337 -0.0510 0.0859 0.0604
3.750 0.7106 0.01231 0.00348 -0.0510 0.0858 0.0670
4.000 0.7375 0.01237 0.00361 -0.0510 0.0856 0.0971
4.250 0.7642 0.01246 0.00374 -0.0510 0.0856 0.1098
4.500 0.7918 0.01236 0.00389 -0.0514 0.0855 0.2376
4.750 0.8232 0.01183 0.00406 -0.0531 0.0854 0.5138
5.000 0.8590 0.01129 0.00432 -0.0557 0.0835 0.7762
5.250 0.8818 0.01141 0.00468 -0.0548 0.0803 0.8505
5.500 0.9047 0.01166 0.00498 -0.0539 0.0782 0.8725
5.750 0.9288 0.01186 0.00522 -0.0533 0.0773 0.8835
6.000 0.9546 0.01207 0.00544 -0.0531 0.0760 0.8918
6.250 0.9770 0.01230 0.00569 -0.0522 0.0671 0.8975
6.500 0.9953 0.01271 0.00606 -0.0504 0.0611 0.9058
6.750 1.0173 0.01307 0.00641 -0.0496 0.0592 0.9110
7.000 1.0358 0.01334 0.00671 -0.0479 0.0582 0.9148
7.250 1.0544 0.01362 0.00703 -0.0462 0.0572 0.9193
7.500 1.0748 0.01394 0.00736 -0.0450 0.0561 0.9239
7.750 1.0957 0.01425 0.00770 -0.0440 0.0550 0.9275
8.000 1.1116 0.01449 0.00800 -0.0419 0.0543 0.9308
8.250 1.1261 0.01476 0.00833 -0.0395 0.0537 0.9356
8.500 1.1424 0.01509 0.00871 -0.0376 0.0527 0.9407
8.750 1.1493 0.01530 0.00899 -0.0337 0.0515 0.9463
9.000 1.1583 0.01564 0.00940 -0.0305 0.0500 0.9509
9.250 1.1699 0.01615 0.00996 -0.0280 0.0480 0.9545
9.500 1.1913 0.01636 0.01020 -0.0274 0.0452 0.9565
9.750 1.2010 0.01660 0.01035 -0.0245 0.0287 0.9609
10.000 1.2060 0.01725 0.01099 -0.0212 0.0222 0.9650
10.250 1.2112 0.01791 0.01169 -0.0181 0.0195 0.9704
10.500 1.2152 0.01859 0.01242 -0.0150 0.0177 0.9758
10.750 1.2203 0.01939 0.01327 -0.0124 0.0164 0.9821
11.000 1.2333 0.02026 0.01421 -0.0116 0.0160 0.9908
11.250 1.2473 0.02133 0.01535 -0.0114 0.0156 1.0000
11.500 1.2605 0.02253 0.01662 -0.0112 0.0152 1.0000
11.750 1.2725 0.02388 0.01804 -0.0111 0.0149 1.0000
12.000 1.2836 0.02536 0.01959 -0.0110 0.0145 1.0000
12.250 1.2933 0.02700 0.02129 -0.0110 0.0141 1.0000
12.500 1.3017 0.02881 0.02317 -0.0110 0.0138 1.0000
12.750 1.3089 0.03076 0.02519 -0.0111 0.0134 1.0000
13.000 1.3147 0.03288 0.02738 -0.0112 0.0129 1.0000
13.250 1.3201 0.03505 0.02962 -0.0113 0.0127 1.0000
13.500 1.3226 0.03756 0.03221 -0.0114 0.0122 1.0000
13.750 1.3237 0.04024 0.03498 -0.0116 0.0118 1.0000
14.000 1.3229 0.04313 0.03798 -0.0118 0.0115 1.0000
14.500 1.3262 0.04863 0.04366 -0.0127 0.0111 1.0000
14.750 1.3355 0.05062 0.04569 -0.0132 0.0109 1.0000
15.000 1.3431 0.05283 0.04796 -0.0138 0.0106 1.0000
15.250 1.3492 0.05526 0.05046 -0.0145 0.0101 1.0000
15.500 1.3526 0.05803 0.05330 -0.0152 0.0098 1.0000
15.750 1.3577 0.06064 0.05596 -0.0159 0.0092 1.0000
16.000 1.3612 0.06350 0.05888 -0.0168 0.0086 1.0000
16.250 1.3609 0.06688 0.06230 -0.0178 0.0077 1.0000
16.500 1.3633 0.06997 0.06545 -0.0188 0.0071 1.0000
16.750 1.3733 0.07210 0.06764 -0.0197 0.0064 1.0000
17.000 1.3761 0.07524 0.07084 -0.0208 0.0059 1.0000
17.250 1.3815 0.07802 0.07363 -0.0219 0.0049 1.0000
17.500 1.3807 0.08172 0.07737 -0.0232 0.0044 1.0000
17.750 1.3795 0.08553 0.08123 -0.0247 0.0038 1.0000
18.000 1.3788 0.08931 0.08505 -0.0262 0.0034 1.0000
18.250 1.3756 0.09357 0.08938 -0.0279 0.0031 1.0000
18.500 1.3719 0.09795 0.09385 -0.0297 0.0031 1.0000
|
Polar data table (+)
Polar graphs
<< Back to FX 63-143 AIRFOIL (fx63143-il)