ULTIMATE (ultimate-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: ULTIMATE (ultimate-il) Reynolds number: 200,000 Max Cl/Cd: 47.82 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ultimate-il-200000.txt Download as CSV file: xf-ultimate-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: ULTIMATE
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.750 -0.8463 0.08776 0.08403 -0.0179 1.0000 0.0386
-12.500 -0.8689 0.08041 0.07659 -0.0217 1.0000 0.0378
-12.250 -0.8937 0.07387 0.06989 -0.0249 1.0000 0.0372
-12.000 -0.9189 0.06811 0.06395 -0.0268 1.0000 0.0366
-11.750 -0.9433 0.06308 0.05871 -0.0271 1.0000 0.0359
-11.500 -0.9686 0.05843 0.05380 -0.0257 1.0000 0.0351
-11.250 -0.9953 0.05414 0.04916 -0.0222 1.0000 0.0342
-11.000 -1.0177 0.05047 0.04511 -0.0174 1.0000 0.0335
-10.750 -1.0211 0.04769 0.04207 -0.0145 1.0000 0.0338
-10.500 -1.0212 0.04483 0.03893 -0.0118 1.0000 0.0341
-10.250 -1.0200 0.04157 0.03528 -0.0088 1.0000 0.0339
-10.000 -1.0137 0.03844 0.03174 -0.0063 1.0000 0.0337
-9.750 -1.0008 0.03588 0.02888 -0.0045 1.0000 0.0340
-9.500 -0.9845 0.03365 0.02642 -0.0031 1.0000 0.0345
-9.250 -0.9659 0.03163 0.02416 -0.0019 1.0000 0.0353
-9.000 -0.9461 0.02992 0.02225 -0.0008 1.0000 0.0365
-8.750 -0.9259 0.02862 0.02076 0.0003 1.0000 0.0382
-8.500 -0.9056 0.02768 0.01959 0.0016 1.0000 0.0395
-8.250 -0.8815 0.02487 0.01677 0.0017 1.0000 0.0414
-8.000 -0.8611 0.02353 0.01544 0.0027 1.0000 0.0434
-7.750 -0.8413 0.02247 0.01432 0.0039 1.0000 0.0460
-7.500 -0.8212 0.02167 0.01343 0.0052 1.0000 0.0490
-7.250 -0.8072 0.02022 0.01200 0.0074 1.0000 0.0525
-7.000 -0.7917 0.01931 0.01109 0.0094 1.0000 0.0565
-6.750 -0.7754 0.01862 0.01033 0.0115 1.0000 0.0612
-6.500 -0.7660 0.01763 0.00939 0.0147 1.0000 0.0686
-6.250 -0.7558 0.01684 0.00863 0.0180 1.0000 0.0790
-6.000 -0.7489 0.01584 0.00784 0.0217 1.0000 0.1086
-5.750 -0.7494 0.01437 0.00711 0.0264 1.0000 0.2196
-5.500 -0.7419 0.01364 0.00679 0.0300 1.0000 0.3043
-5.250 -0.7288 0.01329 0.00661 0.0326 1.0000 0.3550
-5.000 -0.7135 0.01306 0.00644 0.0348 1.0000 0.3934
-4.750 -0.6971 0.01286 0.00629 0.0368 1.0000 0.4244
-4.500 -0.6797 0.01267 0.00617 0.0387 1.0000 0.4539
-4.250 -0.6618 0.01249 0.00608 0.0404 1.0000 0.4813
-4.000 -0.6306 0.01227 0.00598 0.0393 0.9966 0.5194
-3.750 -0.5931 0.01206 0.00597 0.0372 0.9910 0.5666
-3.500 -0.5534 0.01192 0.00593 0.0345 0.9856 0.6026
-3.250 -0.5158 0.01179 0.00586 0.0324 0.9791 0.6330
-3.000 -0.4749 0.01169 0.00582 0.0297 0.9740 0.6638
-2.750 -0.4390 0.01158 0.00577 0.0280 0.9665 0.6922
-2.500 -0.3973 0.01148 0.00572 0.0252 0.9615 0.7204
-2.250 -0.3630 0.01138 0.00567 0.0241 0.9532 0.7465
-2.000 -0.3242 0.01128 0.00561 0.0221 0.9471 0.7701
-1.750 -0.2938 0.01118 0.00555 0.0219 0.9371 0.7921
-1.500 -0.2587 0.01106 0.00548 0.0209 0.9299 0.8108
-1.250 -0.2291 0.01096 0.00540 0.0211 0.9195 0.8285
-1.000 -0.2007 0.01086 0.00532 0.0215 0.9088 0.8457
-0.750 -0.1721 0.01074 0.00521 0.0220 0.8989 0.8622
-0.500 -0.1437 0.01063 0.00510 0.0225 0.8879 0.8782
-0.250 -0.1131 0.01056 0.00503 0.0224 0.8752 0.8931
0.000 -0.0780 0.01053 0.00500 0.0213 0.8626 0.9064
0.250 -0.0393 0.01055 0.00501 0.0194 0.8495 0.9178
0.500 -0.0007 0.01060 0.00504 0.0174 0.8355 0.9285
0.750 0.0452 0.01073 0.00516 0.0139 0.8207 0.9338
1.000 0.0838 0.01083 0.00523 0.0119 0.8045 0.9423
1.250 0.1267 0.01099 0.00537 0.0091 0.7875 0.9467
1.500 0.1645 0.01112 0.00548 0.0072 0.7699 0.9541
1.750 0.2041 0.01127 0.00558 0.0050 0.7520 0.9601
2.000 0.2410 0.01142 0.00568 0.0032 0.7329 0.9684
2.250 0.2800 0.01157 0.00578 0.0009 0.7103 0.9756
2.500 0.3184 0.01174 0.00585 -0.0013 0.6891 0.9830
2.750 0.3583 0.01190 0.00600 -0.0040 0.6679 0.9902
3.000 0.4008 0.01204 0.00612 -0.0073 0.6484 0.9965
3.250 0.4375 0.01209 0.00615 -0.0095 0.6301 1.0000
3.500 0.4599 0.01205 0.00607 -0.0088 0.6088 1.0000
3.750 0.4820 0.01201 0.00596 -0.0080 0.5829 1.0000
4.000 0.5044 0.01199 0.00589 -0.0073 0.5552 1.0000
4.250 0.5269 0.01202 0.00585 -0.0065 0.5252 1.0000
4.500 0.5497 0.01209 0.00587 -0.0057 0.4936 1.0000
4.750 0.5726 0.01223 0.00592 -0.0050 0.4591 1.0000
5.000 0.5953 0.01245 0.00602 -0.0043 0.4123 1.0000
5.250 0.6159 0.01303 0.00621 -0.0034 0.3106 1.0000
5.500 0.6284 0.01537 0.00721 -0.0022 0.0954 1.0000
5.750 0.6474 0.01641 0.00810 -0.0010 0.0716 1.0000
6.000 0.6663 0.01734 0.00900 0.0004 0.0625 1.0000
6.250 0.6859 0.01812 0.00981 0.0016 0.0568 1.0000
6.500 0.7018 0.01951 0.01115 0.0034 0.0526 1.0000
6.750 0.7224 0.02021 0.01193 0.0046 0.0495 1.0000
7.000 0.7421 0.02109 0.01283 0.0059 0.0462 1.0000
7.250 0.7612 0.02231 0.01400 0.0073 0.0439 1.0000
7.500 0.7811 0.02451 0.01618 0.0085 0.0420 1.0000
7.750 0.8028 0.02543 0.01727 0.0096 0.0408 1.0000
8.000 0.8239 0.02647 0.01847 0.0108 0.0387 1.0000
8.250 0.8448 0.02787 0.01999 0.0119 0.0373 1.0000
8.500 0.8654 0.02954 0.02180 0.0130 0.0364 1.0000
8.750 0.8851 0.03144 0.02388 0.0142 0.0357 1.0000
9.000 0.9032 0.03365 0.02633 0.0156 0.0354 1.0000
9.250 0.9168 0.03649 0.02958 0.0177 0.0358 1.0000
9.500 0.9260 0.03962 0.03313 0.0202 0.0362 1.0000
9.750 0.9330 0.04263 0.03649 0.0226 0.0362 1.0000
10.000 0.9315 0.04654 0.04084 0.0258 0.0370 1.0000
10.250 0.9228 0.05155 0.04626 0.0291 0.0394 1.0000
10.500 0.8456 0.04346 0.03866 0.0373 0.0397 1.0000
10.750 0.8273 0.04764 0.04297 0.0410 0.0407 1.0000
11.000 0.7710 0.07759 0.07381 0.0382 0.0745 1.0000
11.250 0.7406 0.08461 0.08090 0.0344 0.0750 1.0000
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Polar data table (+)
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