AG08 (ag08-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: AG08 (ag08-il) Reynolds number: 200,000 Max Cl/Cd: 56.62 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ag08-il-200000-n5.txt Download as CSV file: xf-ag08-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: AG08
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.5921 0.10889 0.10534 0.0193 1.0000 0.0114
-9.000 -0.5894 0.10486 0.10134 0.0178 1.0000 0.0113
-8.750 -0.5868 0.10074 0.09724 0.0160 1.0000 0.0112
-8.500 -0.5843 0.09621 0.09274 0.0134 1.0000 0.0114
-8.000 -0.5761 0.09004 0.08664 0.0103 1.0000 0.0130
-7.750 -0.5717 0.08649 0.08312 0.0079 1.0000 0.0136
-7.500 -0.5659 0.08201 0.07868 0.0036 1.0000 0.0141
-7.250 -0.5571 0.07663 0.07331 -0.0022 1.0000 0.0143
-7.000 -0.5456 0.07037 0.06703 -0.0090 1.0000 0.0140
-6.750 -0.5310 0.06278 0.05938 -0.0171 1.0000 0.0134
-6.500 -0.5128 0.05295 0.04937 -0.0262 1.0000 0.0129
-6.250 -0.4944 0.03536 0.03099 -0.0361 1.0000 0.0120
-6.000 -0.4734 0.02887 0.02378 -0.0383 1.0000 0.0122
-5.750 -0.4499 0.02522 0.01961 -0.0389 1.0000 0.0127
-5.500 -0.4252 0.02254 0.01646 -0.0390 1.0000 0.0134
-5.250 -0.3997 0.02031 0.01380 -0.0389 1.0000 0.0144
-5.000 -0.3735 0.01898 0.01213 -0.0387 1.0000 0.0164
-4.750 -0.3474 0.01736 0.01014 -0.0383 1.0000 0.0178
-4.500 -0.3217 0.01576 0.00836 -0.0380 1.0000 0.0194
-4.250 -0.2955 0.01483 0.00730 -0.0376 1.0000 0.0215
-4.000 -0.2692 0.01415 0.00650 -0.0372 1.0000 0.0249
-3.750 -0.2433 0.01323 0.00551 -0.0368 1.0000 0.0288
-3.500 -0.2171 0.01263 0.00484 -0.0364 1.0000 0.0336
-3.250 -0.1910 0.01209 0.00427 -0.0360 1.0000 0.0414
-3.000 -0.1649 0.01164 0.00376 -0.0356 1.0000 0.0503
-2.750 -0.1390 0.01124 0.00341 -0.0352 1.0000 0.0648
-2.500 -0.1131 0.01089 0.00309 -0.0348 1.0000 0.0816
-2.250 -0.0873 0.01057 0.00283 -0.0344 1.0000 0.1051
-2.000 -0.0615 0.01026 0.00263 -0.0341 1.0000 0.1350
-1.750 -0.0357 0.00995 0.00248 -0.0338 1.0000 0.1775
-1.500 -0.0099 0.00964 0.00238 -0.0336 1.0000 0.2327
-1.250 0.0260 0.00926 0.00227 -0.0356 0.9918 0.3161
-1.000 0.0625 0.00881 0.00219 -0.0376 0.9795 0.4241
-0.750 0.0958 0.00812 0.00212 -0.0388 0.9646 0.5960
-0.500 0.1333 0.00717 0.00206 -0.0397 0.9484 1.0000
-0.250 0.1690 0.00717 0.00194 -0.0411 0.9237 1.0000
0.000 0.2015 0.00719 0.00185 -0.0417 0.8927 1.0000
0.250 0.2307 0.00724 0.00176 -0.0414 0.8568 1.0000
0.500 0.2570 0.00734 0.00170 -0.0406 0.8164 1.0000
0.750 0.2823 0.00749 0.00165 -0.0395 0.7742 1.0000
1.000 0.3075 0.00768 0.00164 -0.0386 0.7307 1.0000
1.250 0.3329 0.00790 0.00165 -0.0377 0.6859 1.0000
1.500 0.3585 0.00815 0.00168 -0.0370 0.6406 1.0000
1.750 0.3843 0.00842 0.00175 -0.0364 0.5961 1.0000
2.000 0.4102 0.00870 0.00183 -0.0358 0.5523 1.0000
2.250 0.4363 0.00900 0.00194 -0.0354 0.5096 1.0000
2.500 0.4626 0.00930 0.00207 -0.0350 0.4695 1.0000
2.750 0.4889 0.00960 0.00225 -0.0347 0.4323 1.0000
3.000 0.5153 0.00992 0.00242 -0.0344 0.3972 1.0000
3.250 0.5418 0.01023 0.00262 -0.0341 0.3643 1.0000
3.500 0.5682 0.01054 0.00284 -0.0338 0.3333 1.0000
3.750 0.5947 0.01088 0.00311 -0.0336 0.3034 1.0000
4.000 0.6211 0.01123 0.00338 -0.0334 0.2744 1.0000
4.250 0.6475 0.01159 0.00367 -0.0331 0.2465 1.0000
4.500 0.6737 0.01197 0.00399 -0.0329 0.2189 1.0000
4.750 0.6999 0.01238 0.00437 -0.0327 0.1915 1.0000
5.000 0.7259 0.01282 0.00475 -0.0324 0.1642 1.0000
5.250 0.7517 0.01332 0.00518 -0.0322 0.1365 1.0000
5.500 0.7772 0.01388 0.00566 -0.0319 0.1089 1.0000
5.750 0.8025 0.01450 0.00620 -0.0317 0.0831 1.0000
6.000 0.8276 0.01518 0.00685 -0.0313 0.0603 1.0000
6.250 0.8524 0.01593 0.00759 -0.0310 0.0432 1.0000
6.500 0.8769 0.01677 0.00844 -0.0305 0.0312 1.0000
6.750 0.9007 0.01778 0.00950 -0.0300 0.0240 1.0000
7.000 0.9247 0.01871 0.01060 -0.0294 0.0203 1.0000
7.250 0.9466 0.02006 0.01202 -0.0288 0.0166 1.0000
7.500 0.9701 0.02102 0.01320 -0.0282 0.0149 1.0000
7.750 0.9922 0.02231 0.01468 -0.0274 0.0136 1.0000
8.000 1.0135 0.02370 0.01624 -0.0266 0.0125 1.0000
8.250 1.0339 0.02528 0.01798 -0.0258 0.0117 1.0000
8.500 1.0517 0.02749 0.02040 -0.0248 0.0111 1.0000
8.750 1.0681 0.03013 0.02334 -0.0236 0.0106 1.0000
9.000 1.0867 0.03191 0.02546 -0.0227 0.0099 1.0000
9.250 1.1032 0.03395 0.02784 -0.0218 0.0090 1.0000
9.500 1.1159 0.03662 0.03092 -0.0207 0.0086 1.0000
9.750 1.1246 0.03975 0.03444 -0.0196 0.0083 1.0000
10.000 1.1289 0.04319 0.03828 -0.0184 0.0081 1.0000
10.250 1.1278 0.04694 0.04239 -0.0174 0.0079 1.0000
10.500 1.1196 0.05093 0.04671 -0.0165 0.0078 1.0000
10.750 1.1035 0.05530 0.05135 -0.0164 0.0078 1.0000
11.000 1.0848 0.06129 0.05760 -0.0200 0.0079 1.0000
11.250 1.0636 0.07023 0.06678 -0.0281 0.0080 1.0000
11.500 1.0357 0.08316 0.07993 -0.0391 0.0083 1.0000
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Polar data table (+)
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