GOE 55 AIRFOIL (goe55-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 55 AIRFOIL (goe55-il) Reynolds number: 500,000 Max Cl/Cd: 64.63 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe55-il-500000-n5.txt Download as CSV file: xf-goe55-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 55 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.6743 0.11070 0.10836 0.0418 1.0000 0.0156
-9.250 -0.6736 0.10608 0.10374 0.0395 1.0000 0.0161
-9.000 -0.6881 0.09731 0.09499 0.0342 1.0000 0.0174
-8.750 -0.6823 0.09444 0.09213 0.0326 1.0000 0.0175
-8.500 -0.6766 0.09159 0.08930 0.0310 1.0000 0.0177
-8.250 -0.6704 0.08818 0.08590 0.0282 1.0000 0.0180
-8.000 -0.6610 0.08405 0.08177 0.0239 1.0000 0.0183
-7.750 -0.6500 0.07909 0.07680 0.0183 1.0000 0.0187
-7.250 -0.6224 0.05499 0.05238 -0.0091 1.0000 0.0214
-7.000 -0.5997 0.05113 0.04842 -0.0126 1.0000 0.0217
-6.750 -0.5759 0.04760 0.04478 -0.0155 1.0000 0.0221
-6.500 -0.5496 0.02193 0.01716 -0.0293 1.0000 0.0256
-6.250 -0.5224 0.01973 0.01467 -0.0300 1.0000 0.0263
-6.000 -0.4947 0.01870 0.01350 -0.0302 1.0000 0.0266
-5.750 -0.4667 0.01773 0.01240 -0.0304 1.0000 0.0270
-5.500 -0.4385 0.01682 0.01136 -0.0306 1.0000 0.0273
-5.250 -0.4102 0.01595 0.01036 -0.0307 1.0000 0.0277
-5.000 -0.3818 0.01513 0.00940 -0.0308 1.0000 0.0282
-4.750 -0.3532 0.01438 0.00853 -0.0309 1.0000 0.0286
-4.500 -0.3247 0.01377 0.00783 -0.0309 1.0000 0.0293
-4.250 -0.2961 0.01318 0.00716 -0.0309 1.0000 0.0299
-4.000 -0.2674 0.01259 0.00648 -0.0309 1.0000 0.0303
-3.750 -0.2373 0.01223 0.00596 -0.0310 0.8981 0.0307
-3.500 -0.2179 0.01273 0.00556 -0.0285 0.6552 0.0309
-3.250 -0.1899 0.01279 0.00510 -0.0285 0.5178 0.0312
-3.000 -0.1613 0.01253 0.00455 -0.0287 0.4469 0.0317
-2.750 -0.1326 0.01221 0.00408 -0.0288 0.4116 0.0323
-2.500 -0.1039 0.01199 0.00376 -0.0288 0.3897 0.0330
-2.250 -0.0752 0.01181 0.00351 -0.0288 0.3742 0.0337
-2.000 -0.0465 0.01167 0.00331 -0.0288 0.3631 0.0346
-1.500 0.0111 0.01138 0.00293 -0.0287 0.3478 0.0367
-1.250 0.0400 0.01125 0.00276 -0.0287 0.3418 0.0375
-1.000 0.0688 0.01116 0.00262 -0.0286 0.3368 0.0383
-0.750 0.0977 0.01099 0.00246 -0.0286 0.3325 0.0402
-0.500 0.1265 0.01090 0.00236 -0.0286 0.3268 0.0423
-0.250 0.1553 0.01086 0.00228 -0.0285 0.3208 0.0451
0.000 0.1841 0.01077 0.00222 -0.0285 0.3154 0.0498
0.250 0.2129 0.01071 0.00218 -0.0284 0.3099 0.0593
0.500 0.2417 0.01066 0.00220 -0.0284 0.3040 0.0811
0.750 0.2704 0.01063 0.00220 -0.0284 0.2960 0.0967
1.000 0.2990 0.01064 0.00218 -0.0283 0.2869 0.1069
1.250 0.3277 0.01061 0.00217 -0.0283 0.2772 0.1178
1.500 0.3566 0.01047 0.00220 -0.0284 0.2703 0.1769
2.000 0.4066 0.00839 0.00224 -0.0271 0.2560 1.0000
2.500 0.4639 0.00861 0.00232 -0.0269 0.2308 1.0000
2.750 0.4923 0.00886 0.00237 -0.0269 0.1908 1.0000
3.000 0.5204 0.00930 0.00256 -0.0270 0.1469 1.0000
3.250 0.5487 0.00962 0.00275 -0.0270 0.1251 1.0000
3.500 0.5771 0.00989 0.00293 -0.0270 0.1096 1.0000
3.750 0.6054 0.01016 0.00313 -0.0269 0.0953 1.0000
4.000 0.6332 0.01080 0.00347 -0.0270 0.0460 1.0000
4.250 0.6614 0.01112 0.00374 -0.0270 0.0401 1.0000
4.500 0.6896 0.01141 0.00403 -0.0269 0.0371 1.0000
4.750 0.7178 0.01167 0.00429 -0.0268 0.0359 1.0000
5.000 0.7459 0.01194 0.00458 -0.0267 0.0351 1.0000
5.250 0.7739 0.01223 0.00490 -0.0266 0.0344 1.0000
5.500 0.8018 0.01254 0.00524 -0.0266 0.0336 1.0000
5.750 0.8296 0.01288 0.00561 -0.0265 0.0328 1.0000
6.000 0.8573 0.01327 0.00603 -0.0264 0.0319 1.0000
6.250 0.8848 0.01369 0.00648 -0.0263 0.0312 1.0000
6.500 0.9121 0.01416 0.00699 -0.0262 0.0306 1.0000
6.750 0.9392 0.01471 0.00760 -0.0261 0.0300 1.0000
7.000 0.9659 0.01535 0.00830 -0.0260 0.0295 1.0000
7.250 0.9925 0.01598 0.00898 -0.0259 0.0291 1.0000
7.500 1.0191 0.01649 0.00955 -0.0257 0.0289 1.0000
7.750 1.0455 0.01703 0.01016 -0.0255 0.0286 1.0000
8.000 1.0717 0.01757 0.01077 -0.0253 0.0283 1.0000
8.250 1.0978 0.01814 0.01141 -0.0251 0.0278 1.0000
8.500 1.1234 0.01879 0.01213 -0.0249 0.0274 1.0000
8.750 1.1487 0.01948 0.01291 -0.0247 0.0270 1.0000
9.000 1.1736 0.02020 0.01371 -0.0244 0.0266 1.0000
9.250 1.1984 0.02091 0.01450 -0.0241 0.0260 1.0000
9.500 1.2228 0.02163 0.01531 -0.0238 0.0255 1.0000
9.750 1.2467 0.02244 0.01619 -0.0235 0.0250 1.0000
10.000 1.2701 0.02328 0.01710 -0.0232 0.0246 1.0000
10.250 1.2926 0.02426 0.01814 -0.0228 0.0240 1.0000
10.500 1.3120 0.02583 0.01980 -0.0223 0.0234 1.0000
10.750 1.3336 0.02686 0.02096 -0.0218 0.0230 1.0000
11.000 1.3560 0.02762 0.02186 -0.0214 0.0227 1.0000
11.250 1.3776 0.02849 0.02287 -0.0209 0.0222 1.0000
11.500 1.3980 0.02948 0.02400 -0.0204 0.0217 1.0000
11.750 1.4176 0.03054 0.02520 -0.0199 0.0211 1.0000
12.000 1.4366 0.03156 0.02635 -0.0193 0.0206 1.0000
12.250 1.4550 0.03258 0.02748 -0.0188 0.0201 1.0000
12.500 1.4726 0.03363 0.02862 -0.0182 0.0197 1.0000
12.750 1.4886 0.03480 0.02988 -0.0176 0.0193 1.0000
13.000 1.5025 0.03611 0.03130 -0.0170 0.0190 1.0000
13.250 1.5114 0.03790 0.03318 -0.0162 0.0186 1.0000
13.500 1.5158 0.03998 0.03542 -0.0153 0.0184 1.0000
13.750 1.5117 0.04255 0.03819 -0.0143 0.0181 1.0000
14.000 1.5031 0.04708 0.04294 -0.0171 0.0179 1.0000
14.250 1.4936 0.05435 0.05045 -0.0243 0.0177 1.0000
14.500 1.4797 0.06218 0.05847 -0.0306 0.0176 1.0000
14.750 1.4597 0.07008 0.06653 -0.0358 0.0175 1.0000
15.000 1.4364 0.07806 0.07465 -0.0404 0.0174 1.0000
15.250 1.4116 0.08613 0.08284 -0.0448 0.0174 1.0000
15.500 1.3852 0.09450 0.09134 -0.0493 0.0174 1.0000
15.750 1.3585 0.10309 0.10006 -0.0538 0.0173 1.0000
16.000 1.3316 0.11230 0.10940 -0.0588 0.0173 1.0000
16.250 1.3032 0.12249 0.11973 -0.0644 0.0172 1.0000
16.500 1.2681 0.13489 0.13229 -0.0714 0.0172 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 55 AIRFOIL (goe55-il)