OAF128 AIRFOIL (oaf128-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: OAF128 AIRFOIL (oaf128-il) Reynolds number: 1,000,000 Max Cl/Cd: 85.39 at α=8.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-oaf128-il-1000000.txt Download as CSV file: xf-oaf128-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: OAF128 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
------ -------- --------- --------- -------- -------- --------
-16.500 -1.3157 0.08341 0.07994 0.0150 1.0000 0.0173
-16.250 -1.3327 0.07718 0.07362 0.0118 1.0000 0.0173
-16.000 -1.3499 0.07099 0.06732 0.0086 1.0000 0.0174
-15.750 -1.3665 0.06494 0.06117 0.0055 1.0000 0.0174
-15.500 -1.3818 0.05901 0.05513 0.0022 1.0000 0.0175
-15.250 -1.3949 0.05332 0.04933 -0.0010 1.0000 0.0175
-15.000 -1.4049 0.04794 0.04383 -0.0044 1.0000 0.0176
-14.750 -1.4113 0.04286 0.03863 -0.0079 1.0000 0.0177
-14.500 -1.4137 0.03821 0.03385 -0.0117 1.0000 0.0178
-14.250 -1.4163 0.03402 0.02952 -0.0151 1.0000 0.0179
-14.000 -1.4138 0.03122 0.02659 -0.0168 1.0000 0.0180
-13.750 -1.4009 0.02958 0.02485 -0.0171 1.0000 0.0182
-13.500 -1.3842 0.02826 0.02344 -0.0172 1.0000 0.0184
-13.250 -1.3649 0.02715 0.02225 -0.0172 1.0000 0.0186
-13.000 -1.3439 0.02620 0.02122 -0.0172 1.0000 0.0188
-12.750 -1.3279 0.02437 0.01929 -0.0174 1.0000 0.0192
-12.500 -1.3072 0.02310 0.01796 -0.0175 1.0000 0.0195
-12.250 -1.2838 0.02216 0.01698 -0.0176 1.0000 0.0198
-12.000 -1.2593 0.02137 0.01615 -0.0178 1.0000 0.0202
-11.750 -1.2341 0.02064 0.01537 -0.0179 1.0000 0.0205
-11.500 -1.2083 0.01995 0.01464 -0.0180 1.0000 0.0209
-11.250 -1.1820 0.01931 0.01394 -0.0182 1.0000 0.0213
-11.000 -1.1552 0.01872 0.01330 -0.0184 1.0000 0.0217
-10.750 -1.1279 0.01822 0.01275 -0.0185 1.0000 0.0220
-10.500 -1.1018 0.01721 0.01168 -0.0189 1.0000 0.0227
-10.250 -1.0742 0.01658 0.01104 -0.0192 1.0000 0.0233
-10.000 -1.0462 0.01608 0.01051 -0.0195 1.0000 0.0239
-9.750 -1.0178 0.01561 0.01001 -0.0197 1.0000 0.0246
-9.500 -0.9892 0.01520 0.00956 -0.0200 1.0000 0.0252
-9.250 -0.9606 0.01466 0.00899 -0.0203 1.0000 0.0261
-9.000 -0.9318 0.01413 0.00845 -0.0207 1.0000 0.0271
-8.750 -0.9027 0.01373 0.00804 -0.0210 1.0000 0.0281
-8.500 -0.8734 0.01338 0.00767 -0.0212 1.0000 0.0291
-8.250 -0.8440 0.01289 0.00716 -0.0216 1.0000 0.0305
-8.000 -0.8145 0.01250 0.00677 -0.0220 1.0000 0.0321
-7.750 -0.7848 0.01219 0.00645 -0.0223 1.0000 0.0337
-7.500 -0.7549 0.01174 0.00602 -0.0227 1.0000 0.0365
-7.250 -0.7250 0.01137 0.00567 -0.0231 1.0000 0.0399
-7.000 -0.6949 0.01098 0.00531 -0.0236 1.0000 0.0455
-6.750 -0.6646 0.01054 0.00495 -0.0241 1.0000 0.0560
-6.500 -0.6343 0.01008 0.00462 -0.0247 1.0000 0.0734
-6.250 -0.6039 0.00975 0.00437 -0.0252 1.0000 0.0881
-6.000 -0.5747 0.00965 0.00429 -0.0252 0.9468 0.0992
-5.750 -0.5494 0.00956 0.00417 -0.0243 0.9117 0.1090
-5.500 -0.5220 0.00944 0.00401 -0.0240 0.8852 0.1180
-5.250 -0.4936 0.00934 0.00385 -0.0238 0.8588 0.1266
-5.000 -0.4646 0.00921 0.00368 -0.0238 0.8340 0.1366
-4.750 -0.4351 0.00905 0.00351 -0.0239 0.8095 0.1506
-4.500 -0.4054 0.00891 0.00336 -0.0241 0.7847 0.1678
-4.250 -0.3756 0.00882 0.00323 -0.0243 0.7597 0.1826
-4.000 -0.3456 0.00874 0.00311 -0.0245 0.7349 0.1952
-3.750 -0.3155 0.00863 0.00298 -0.0248 0.7117 0.2078
-3.500 -0.2852 0.00853 0.00286 -0.0251 0.6895 0.2230
-3.250 -0.2550 0.00842 0.00276 -0.0254 0.6681 0.2441
-2.750 -0.1944 0.00830 0.00260 -0.0260 0.6233 0.2795
-2.500 -0.1641 0.00825 0.00253 -0.0263 0.6030 0.2961
-2.250 -0.1337 0.00821 0.00248 -0.0266 0.5860 0.3120
-2.000 -0.1034 0.00818 0.00244 -0.0269 0.5685 0.3258
-1.750 -0.0731 0.00817 0.00239 -0.0271 0.5512 0.3389
-1.250 -0.0125 0.00813 0.00234 -0.0276 0.5247 0.3637
-1.000 0.0178 0.00814 0.00232 -0.0279 0.5123 0.3748
-0.750 0.0481 0.00814 0.00231 -0.0281 0.4999 0.3867
-0.500 0.0784 0.00813 0.00231 -0.0284 0.4881 0.3990
-0.250 0.1087 0.00814 0.00231 -0.0286 0.4765 0.4114
0.250 0.1693 0.00814 0.00233 -0.0291 0.4561 0.4403
0.500 0.1996 0.00813 0.00234 -0.0294 0.4456 0.4536
0.750 0.2298 0.00816 0.00236 -0.0296 0.4352 0.4663
1.000 0.2601 0.00818 0.00239 -0.0298 0.4251 0.4804
1.250 0.2903 0.00818 0.00242 -0.0301 0.4157 0.4954
1.500 0.3206 0.00823 0.00246 -0.0304 0.4029 0.5105
1.750 0.3508 0.00827 0.00249 -0.0306 0.3871 0.5265
2.000 0.3810 0.00831 0.00254 -0.0309 0.3719 0.5447
2.250 0.4112 0.00833 0.00260 -0.0311 0.3594 0.5663
2.500 0.4414 0.00838 0.00266 -0.0314 0.3470 0.5896
2.750 0.4715 0.00843 0.00274 -0.0317 0.3333 0.6148
3.000 0.5017 0.00846 0.00282 -0.0320 0.3200 0.6448
3.250 0.5317 0.00848 0.00292 -0.0322 0.3064 0.6829
3.500 0.5616 0.00853 0.00304 -0.0325 0.2915 0.7269
3.750 0.5912 0.00858 0.00317 -0.0326 0.2767 0.7683
4.000 0.6205 0.00866 0.00330 -0.0327 0.2610 0.8082
4.250 0.6490 0.00875 0.00344 -0.0326 0.2452 0.8492
4.500 0.6762 0.00882 0.00357 -0.0322 0.2316 0.8934
4.750 0.6987 0.00865 0.00351 -0.0306 0.2209 1.0000
5.000 0.7288 0.00886 0.00366 -0.0309 0.2091 1.0000
5.250 0.7587 0.00909 0.00383 -0.0312 0.1970 1.0000
5.500 0.7886 0.00933 0.00400 -0.0315 0.1824 1.0000
5.750 0.8183 0.00962 0.00419 -0.0318 0.1635 1.0000
6.000 0.8478 0.01002 0.00444 -0.0322 0.1408 1.0000
6.250 0.8772 0.01040 0.00472 -0.0325 0.1222 1.0000
6.500 0.9064 0.01077 0.00500 -0.0328 0.1086 1.0000
6.750 0.9357 0.01111 0.00528 -0.0331 0.0996 1.0000
7.000 0.9648 0.01146 0.00558 -0.0333 0.0920 1.0000
7.250 0.9939 0.01174 0.00584 -0.0335 0.0874 1.0000
7.500 1.0227 0.01211 0.00616 -0.0337 0.0821 1.0000
7.750 1.0517 0.01235 0.00642 -0.0339 0.0791 1.0000
8.000 1.0803 0.01270 0.00674 -0.0340 0.0754 1.0000
8.250 1.1089 0.01304 0.00708 -0.0342 0.0722 1.0000
8.500 1.1374 0.01332 0.00737 -0.0343 0.0699 1.0000
8.750 1.1656 0.01368 0.00771 -0.0344 0.0670 1.0000
9.000 1.1936 0.01408 0.00811 -0.0345 0.0644 1.0000
9.250 1.2217 0.01436 0.00842 -0.0346 0.0627 1.0000
9.500 1.2495 0.01471 0.00876 -0.0346 0.0605 1.0000
9.750 1.2765 0.01522 0.00926 -0.0346 0.0579 1.0000
10.000 1.3041 0.01554 0.00962 -0.0346 0.0566 1.0000
10.250 1.3315 0.01588 0.00997 -0.0346 0.0550 1.0000
10.500 1.3584 0.01629 0.01038 -0.0346 0.0532 1.0000
10.750 1.3840 0.01692 0.01101 -0.0345 0.0510 1.0000
11.000 1.4109 0.01724 0.01137 -0.0344 0.0500 1.0000
11.250 1.4373 0.01762 0.01178 -0.0343 0.0486 1.0000
11.500 1.4630 0.01808 0.01225 -0.0342 0.0471 1.0000
11.750 1.4873 0.01876 0.01293 -0.0340 0.0454 1.0000
12.000 1.5125 0.01922 0.01344 -0.0337 0.0444 1.0000
12.250 1.5376 0.01965 0.01391 -0.0335 0.0432 1.0000
12.500 1.5618 0.02018 0.01446 -0.0332 0.0419 1.0000
12.750 1.5845 0.02086 0.01515 -0.0328 0.0406 1.0000
13.000 1.6064 0.02159 0.01593 -0.0323 0.0394 1.0000
13.250 1.6295 0.02210 0.01649 -0.0319 0.0385 1.0000
13.500 1.6515 0.02270 0.01713 -0.0314 0.0374 1.0000
13.750 1.6718 0.02344 0.01789 -0.0308 0.0364 1.0000
14.000 1.6881 0.02450 0.01897 -0.0298 0.0352 1.0000
14.250 1.7067 0.02524 0.01978 -0.0290 0.0344 1.0000
14.500 1.7236 0.02603 0.02064 -0.0280 0.0337 1.0000
14.750 1.7376 0.02696 0.02164 -0.0268 0.0329 1.0000
15.000 1.7468 0.02812 0.02285 -0.0252 0.0322 1.0000
15.250 1.7465 0.02993 0.02473 -0.0234 0.0316 1.0000
15.500 1.7441 0.03259 0.02748 -0.0230 0.0309 1.0000
15.750 1.7475 0.03509 0.03008 -0.0233 0.0304 1.0000
16.000 1.7522 0.03762 0.03271 -0.0239 0.0299 1.0000
16.250 1.7542 0.04065 0.03584 -0.0249 0.0294 1.0000
16.500 1.7530 0.04418 0.03948 -0.0261 0.0289 1.0000
16.750 1.7480 0.04830 0.04369 -0.0277 0.0284 1.0000
17.000 1.7389 0.05298 0.04848 -0.0294 0.0281 1.0000
17.250 1.7252 0.05831 0.05391 -0.0314 0.0278 1.0000
17.500 1.7070 0.06426 0.05998 -0.0337 0.0275 1.0000
17.750 1.6846 0.07084 0.06669 -0.0362 0.0272 1.0000
18.000 1.6590 0.07795 0.07392 -0.0390 0.0270 1.0000
18.250 1.6331 0.08535 0.08144 -0.0421 0.0268 1.0000
|
Polar data table (+)
Polar graphs
<< Back to OAF128 AIRFOIL (oaf128-il)