NACA 0010-66 (naca001066-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA 0010-66 (naca001066-il) Reynolds number: 100,000 Max Cl/Cd: 28.88 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-naca001066-il-100000-n5.txt Download as CSV file: xf-naca001066-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 0010-66
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.7094 0.09433 0.08914 -0.0338 1.0000 0.0341
-11.000 -0.7356 0.08528 0.08013 -0.0387 1.0000 0.0337
-10.750 -0.7659 0.07875 0.07354 -0.0393 1.0000 0.0334
-10.500 -0.7967 0.07391 0.06861 -0.0368 1.0000 0.0332
-10.250 -0.8274 0.07014 0.06473 -0.0321 1.0000 0.0330
-10.000 -0.8579 0.06711 0.06158 -0.0257 1.0000 0.0330
-9.750 -0.8857 0.06381 0.05810 -0.0190 1.0000 0.0330
-9.500 -0.9092 0.05994 0.05397 -0.0127 1.0000 0.0331
-9.250 -0.9294 0.05595 0.04964 -0.0064 1.0000 0.0335
-9.000 -0.9461 0.05200 0.04523 0.0001 1.0000 0.0343
-8.750 -0.9586 0.04842 0.04103 0.0064 1.0000 0.0353
-8.500 -0.9579 0.04588 0.03833 0.0103 1.0000 0.0364
-8.250 -0.9513 0.04426 0.03657 0.0135 1.0000 0.0376
-8.000 -0.9450 0.04223 0.03427 0.0170 1.0000 0.0388
-7.750 -0.9369 0.04017 0.03189 0.0203 1.0000 0.0402
-7.500 -0.9272 0.03824 0.02954 0.0235 1.0000 0.0425
-7.250 -0.9155 0.03621 0.02697 0.0267 1.0000 0.0446
-7.000 -0.9003 0.03419 0.02451 0.0291 1.0000 0.0460
-6.750 -0.8829 0.03272 0.02298 0.0307 1.0000 0.0481
-6.500 -0.8650 0.03173 0.02185 0.0324 1.0000 0.0505
-6.250 -0.8448 0.03049 0.02035 0.0338 1.0000 0.0528
-6.000 -0.8225 0.02919 0.01877 0.0350 1.0000 0.0546
-5.750 -0.8002 0.02823 0.01752 0.0362 1.0000 0.0570
-5.500 -0.7780 0.02709 0.01636 0.0370 1.0000 0.0597
-5.250 -0.7560 0.02623 0.01545 0.0380 1.0000 0.0620
-5.000 -0.7337 0.02540 0.01454 0.0391 1.0000 0.0643
-4.750 -0.7117 0.02465 0.01366 0.0402 1.0000 0.0669
-4.500 -0.6902 0.02398 0.01286 0.0415 1.0000 0.0697
-4.250 -0.6702 0.02329 0.01221 0.0429 1.0000 0.0739
-4.000 -0.6497 0.02278 0.01167 0.0443 1.0000 0.0801
-3.750 -0.6294 0.02221 0.01108 0.0458 1.0000 0.0870
-3.500 -0.6086 0.02171 0.01054 0.0471 1.0000 0.0970
-3.250 -0.5876 0.02119 0.01009 0.0484 1.0000 0.1124
-3.000 -0.5662 0.02068 0.00971 0.0495 1.0000 0.1366
-2.750 -0.3803 0.02029 0.01279 0.0172 1.0000 0.8275
-2.500 -0.3835 0.02051 0.01290 0.0242 1.0000 0.8931
-2.250 -0.3473 0.02152 0.01373 0.0233 1.0000 0.9328
-2.000 -0.2870 0.02249 0.01447 0.0169 1.0000 0.9571
-1.750 -0.2070 0.02324 0.01501 0.0058 1.0000 0.9789
-1.500 -0.1050 0.02350 0.01506 -0.0105 1.0000 1.0000
-1.250 -0.0876 0.02343 0.01493 -0.0088 1.0000 1.0000
-1.000 -0.0701 0.02337 0.01482 -0.0070 1.0000 1.0000
-0.750 -0.0526 0.02333 0.01474 -0.0053 1.0000 1.0000
-0.500 -0.0351 0.02330 0.01469 -0.0035 1.0000 1.0000
-0.250 -0.0176 0.02328 0.01465 -0.0018 1.0000 1.0000
0.000 0.0000 0.02327 0.01464 0.0000 1.0000 1.0000
0.250 0.0176 0.02328 0.01465 0.0018 1.0000 1.0000
0.500 0.0351 0.02330 0.01468 0.0035 1.0000 1.0000
0.750 0.0526 0.02333 0.01474 0.0053 1.0000 1.0000
1.000 0.0701 0.02337 0.01482 0.0070 1.0000 1.0000
1.250 0.0876 0.02342 0.01492 0.0088 1.0000 1.0000
1.500 0.1050 0.02349 0.01505 0.0105 1.0000 1.0000
1.750 0.2066 0.02324 0.01501 -0.0057 0.9791 1.0000
2.000 0.2870 0.02248 0.01446 -0.0169 0.9571 1.0000
2.250 0.3473 0.02151 0.01372 -0.0233 0.9328 1.0000
2.500 0.3836 0.02050 0.01290 -0.0242 0.8932 1.0000
2.750 0.3803 0.02028 0.01277 -0.0172 0.8270 1.0000
3.000 0.5661 0.02068 0.00971 -0.0495 0.1366 1.0000
3.250 0.5875 0.02119 0.01007 -0.0484 0.1124 1.0000
3.500 0.6085 0.02170 0.01054 -0.0471 0.0970 1.0000
3.750 0.6293 0.02221 0.01107 -0.0457 0.0870 1.0000
4.000 0.6496 0.02278 0.01167 -0.0443 0.0802 1.0000
4.250 0.6701 0.02328 0.01221 -0.0429 0.0739 1.0000
4.500 0.6902 0.02398 0.01286 -0.0415 0.0697 1.0000
4.750 0.7116 0.02464 0.01365 -0.0402 0.0669 1.0000
5.000 0.7336 0.02540 0.01453 -0.0391 0.0643 1.0000
5.250 0.7559 0.02622 0.01544 -0.0380 0.0620 1.0000
5.500 0.7779 0.02708 0.01636 -0.0370 0.0597 1.0000
5.750 0.8001 0.02823 0.01752 -0.0361 0.0570 1.0000
6.000 0.8224 0.02919 0.01876 -0.0350 0.0546 1.0000
6.250 0.8447 0.03048 0.02034 -0.0338 0.0528 1.0000
6.500 0.8649 0.03172 0.02184 -0.0324 0.0505 1.0000
6.750 0.8828 0.03270 0.02296 -0.0307 0.0480 1.0000
7.000 0.9002 0.03417 0.02450 -0.0291 0.0460 1.0000
7.250 0.9155 0.03620 0.02696 -0.0267 0.0446 1.0000
7.500 0.9271 0.03823 0.02953 -0.0235 0.0425 1.0000
7.750 0.9369 0.04017 0.03189 -0.0203 0.0402 1.0000
8.000 0.9449 0.04222 0.03426 -0.0170 0.0388 1.0000
8.250 0.9513 0.04424 0.03654 -0.0135 0.0376 1.0000
8.500 0.9579 0.04586 0.03831 -0.0103 0.0364 1.0000
8.750 0.9586 0.04840 0.04101 -0.0064 0.0353 1.0000
9.000 0.9462 0.05199 0.04521 -0.0001 0.0343 1.0000
9.250 0.9293 0.05595 0.04964 0.0064 0.0335 1.0000
9.500 0.9094 0.05992 0.05395 0.0127 0.0331 1.0000
9.750 0.8858 0.06381 0.05810 0.0190 0.0330 1.0000
10.000 0.8581 0.06712 0.06158 0.0257 0.0330 1.0000
10.250 0.8275 0.07015 0.06474 0.0321 0.0330 1.0000
10.500 0.7969 0.07393 0.06863 0.0368 0.0332 1.0000
10.750 0.7661 0.07877 0.07356 0.0392 0.0334 1.0000
11.000 0.7359 0.08532 0.08017 0.0386 0.0337 1.0000
11.250 0.7096 0.09438 0.08918 0.0337 0.0341 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 0010-66 (naca001066-il)