B-29 TIP AIRFOIL (b29tip-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: B-29 TIP AIRFOIL (b29tip-il) Reynolds number: 200,000 Max Cl/Cd: 66.42 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-b29tip-il-200000-n5.txt Download as CSV file: xf-b29tip-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: B-29 TIP AIRFOIL
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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4240 0.08641 0.08327 -0.0067 1.0000 0.0327
-8.750 -0.4374 0.08117 0.07809 -0.0127 1.0000 0.0360
-8.250 -0.5263 0.07838 0.07510 -0.0157 1.0000 0.0252
-8.000 -0.5301 0.07442 0.07116 -0.0172 1.0000 0.0245
-7.750 -0.5316 0.06965 0.06637 -0.0197 1.0000 0.0237
-7.500 -0.5316 0.06450 0.06115 -0.0220 1.0000 0.0228
-7.250 -0.5299 0.05895 0.05549 -0.0238 1.0000 0.0221
-7.000 -0.5260 0.05324 0.04960 -0.0249 1.0000 0.0215
-6.750 -0.5169 0.04904 0.04522 -0.0251 1.0000 0.0220
-6.500 -0.4995 0.04422 0.04011 -0.0267 0.9539 0.0226
-6.250 -0.4790 0.03843 0.03382 -0.0281 0.9165 0.0220
-6.000 -0.4650 0.03289 0.02766 -0.0269 0.8899 0.0211
-5.750 -0.4506 0.02848 0.02257 -0.0249 0.8685 0.0206
-5.500 -0.4321 0.02579 0.01938 -0.0232 0.8509 0.0206
-5.250 -0.4116 0.02372 0.01689 -0.0219 0.8352 0.0207
-5.000 -0.3894 0.02199 0.01476 -0.0207 0.8209 0.0209
-4.750 -0.3660 0.02057 0.01303 -0.0197 0.8076 0.0213
-4.500 -0.3418 0.01934 0.01152 -0.0189 0.7950 0.0218
-4.250 -0.3170 0.01826 0.01020 -0.0181 0.7833 0.0223
-4.000 -0.2919 0.01750 0.00926 -0.0175 0.7722 0.0236
-3.750 -0.2666 0.01677 0.00833 -0.0168 0.7616 0.0247
-3.500 -0.2413 0.01598 0.00740 -0.0161 0.7511 0.0254
-3.250 -0.2161 0.01525 0.00655 -0.0154 0.7411 0.0259
-3.000 -0.1912 0.01466 0.00584 -0.0147 0.7318 0.0265
-2.750 -0.1671 0.01396 0.00505 -0.0139 0.7222 0.0279
-2.500 -0.1421 0.01349 0.00452 -0.0133 0.7130 0.0300
-2.250 -0.1167 0.01316 0.00408 -0.0127 0.7045 0.0330
-2.000 -0.0911 0.01281 0.00370 -0.0121 0.6953 0.0392
-1.750 -0.0658 0.01246 0.00338 -0.0116 0.6871 0.0612
-1.500 -0.0408 0.01210 0.00313 -0.0110 0.6786 0.1006
-1.250 -0.0174 0.01154 0.00293 -0.0104 0.6704 0.1963
-1.000 -0.0080 0.00980 0.00274 -0.0072 0.6632 0.5899
-0.750 0.0710 0.00934 0.00314 -0.0163 0.6543 0.9365
-0.500 0.1146 0.00957 0.00320 -0.0192 0.6466 0.9634
-0.250 0.2064 0.00973 0.00315 -0.0324 0.6364 0.9976
0.000 0.2373 0.00971 0.00302 -0.0332 0.6287 1.0000
0.250 0.2615 0.00972 0.00294 -0.0325 0.6210 1.0000
0.500 0.2859 0.00973 0.00287 -0.0319 0.6136 1.0000
0.750 0.3103 0.00975 0.00283 -0.0313 0.6063 1.0000
1.000 0.3349 0.00978 0.00280 -0.0307 0.5992 1.0000
1.250 0.3595 0.00982 0.00279 -0.0301 0.5920 1.0000
1.500 0.3842 0.00986 0.00279 -0.0295 0.5853 1.0000
1.750 0.4090 0.00991 0.00281 -0.0290 0.5784 1.0000
2.000 0.4337 0.00998 0.00286 -0.0284 0.5720 1.0000
2.250 0.4586 0.01004 0.00292 -0.0278 0.5649 1.0000
2.500 0.4833 0.01013 0.00298 -0.0272 0.5590 1.0000
2.750 0.5082 0.01021 0.00309 -0.0267 0.5519 1.0000
3.000 0.5329 0.01032 0.00319 -0.0261 0.5459 1.0000
3.250 0.5578 0.01041 0.00334 -0.0256 0.5389 1.0000
3.500 0.5825 0.01053 0.00346 -0.0250 0.5325 1.0000
3.750 0.6074 0.01064 0.00366 -0.0244 0.5256 1.0000
4.000 0.6320 0.01076 0.00380 -0.0238 0.5176 1.0000
4.250 0.6560 0.01086 0.00391 -0.0230 0.5030 1.0000
4.500 0.6797 0.01094 0.00400 -0.0222 0.4806 1.0000
4.750 0.7031 0.01108 0.00410 -0.0213 0.4578 1.0000
5.000 0.7271 0.01123 0.00432 -0.0206 0.4403 1.0000
5.250 0.7510 0.01140 0.00454 -0.0198 0.4222 1.0000
5.500 0.7738 0.01165 0.00475 -0.0190 0.3908 1.0000
5.750 0.7936 0.01219 0.00502 -0.0178 0.3227 1.0000
6.000 0.8060 0.01369 0.00577 -0.0160 0.1824 1.0000
6.250 0.8213 0.01492 0.00661 -0.0146 0.1099 1.0000
6.500 0.8391 0.01580 0.00731 -0.0134 0.0730 1.0000
6.750 0.8571 0.01662 0.00802 -0.0121 0.0486 1.0000
7.000 0.8744 0.01749 0.00882 -0.0106 0.0321 1.0000
7.250 0.8919 0.01830 0.00970 -0.0091 0.0251 1.0000
7.500 0.9085 0.01918 0.01064 -0.0076 0.0214 1.0000
7.750 0.9255 0.01995 0.01149 -0.0061 0.0189 1.0000
8.000 0.9399 0.02092 0.01253 -0.0043 0.0172 1.0000
8.250 0.9516 0.02210 0.01380 -0.0022 0.0163 1.0000
8.500 0.9646 0.02312 0.01494 -0.0002 0.0156 1.0000
8.750 0.9762 0.02425 0.01618 0.0019 0.0150 1.0000
9.000 0.9869 0.02544 0.01750 0.0040 0.0144 1.0000
9.250 0.9976 0.02663 0.01879 0.0061 0.0138 1.0000
9.500 1.0082 0.02777 0.02001 0.0080 0.0131 1.0000
9.750 1.0158 0.02890 0.02119 0.0103 0.0124 1.0000
10.000 1.0200 0.03055 0.02289 0.0128 0.0117 1.0000
10.250 1.0277 0.03230 0.02474 0.0148 0.0115 1.0000
10.500 1.0368 0.03390 0.02652 0.0165 0.0112 1.0000
10.750 1.0454 0.03570 0.02850 0.0180 0.0111 1.0000
11.000 1.0523 0.03774 0.03079 0.0196 0.0108 1.0000
11.250 1.0573 0.03997 0.03325 0.0210 0.0106 1.0000
11.500 1.0594 0.04248 0.03600 0.0222 0.0105 1.0000
11.750 1.0588 0.04525 0.03901 0.0232 0.0104 1.0000
12.000 1.0552 0.04831 0.04233 0.0239 0.0103 1.0000
12.250 1.0485 0.05178 0.04605 0.0242 0.0102 1.0000
12.500 1.0393 0.05563 0.05015 0.0240 0.0102 1.0000
12.750 1.0281 0.05984 0.05459 0.0233 0.0101 1.0000
13.000 1.0145 0.06460 0.05958 0.0219 0.0101 1.0000
13.250 0.9995 0.06979 0.06497 0.0200 0.0102 1.0000
13.500 0.9819 0.07570 0.07109 0.0173 0.0102 1.0000
13.750 0.9635 0.08215 0.07773 0.0140 0.0102 1.0000
14.000 0.9456 0.08896 0.08470 0.0102 0.0103 1.0000
14.250 0.9234 0.09746 0.09337 0.0052 0.0104 1.0000
14.500 0.9011 0.10691 0.10296 -0.0006 0.0105 1.0000
14.750 0.8753 0.11855 0.11472 -0.0076 0.0106 1.0000
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Polar data table (+)
Polar graphs
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