B-29 TIP AIRFOIL (b29tip-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: B-29 TIP AIRFOIL (b29tip-il) Reynolds number: 100,000 Max Cl/Cd: 50.85 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-b29tip-il-100000-n5.txt Download as CSV file: xf-b29tip-il-100000-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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.5041 0.10289 0.09808 -0.0017 1.0000 0.0558
-9.000 -0.5039 0.09918 0.09442 -0.0043 1.0000 0.0578
-8.750 -0.5106 0.09533 0.09065 -0.0093 1.0000 0.0599
-8.500 -0.4296 0.08208 0.07779 -0.0135 1.0000 0.0633
-8.000 -0.4394 0.07292 0.06871 -0.0185 1.0000 0.0634
-7.500 -0.5133 0.07207 0.06746 -0.0211 1.0000 0.0452
-7.250 -0.5181 0.06304 0.05809 -0.0258 1.0000 0.0360
-7.000 -0.5111 0.05928 0.05429 -0.0258 1.0000 0.0352
-6.750 -0.5033 0.05537 0.05028 -0.0259 1.0000 0.0345
-6.500 -0.4946 0.05133 0.04606 -0.0259 1.0000 0.0338
-6.250 -0.4846 0.04736 0.04187 -0.0255 1.0000 0.0331
-6.000 -0.4737 0.04352 0.03777 -0.0247 1.0000 0.0326
-5.750 -0.4625 0.03999 0.03394 -0.0234 1.0000 0.0321
-5.500 -0.4392 0.03606 0.02956 -0.0242 0.9733 0.0318
-5.250 -0.4092 0.03236 0.02526 -0.0257 0.9482 0.0317
-5.000 -0.3792 0.02959 0.02196 -0.0266 0.9273 0.0326
-4.750 -0.3501 0.02740 0.01924 -0.0269 0.9087 0.0344
-4.500 -0.3226 0.02543 0.01680 -0.0266 0.8916 0.0350
-4.250 -0.2964 0.02373 0.01470 -0.0259 0.8754 0.0354
-4.000 -0.2704 0.02231 0.01295 -0.0252 0.8602 0.0359
-3.750 -0.2446 0.02113 0.01148 -0.0244 0.8462 0.0365
-3.500 -0.2198 0.01979 0.01004 -0.0236 0.8333 0.0377
-3.250 -0.1955 0.01888 0.00906 -0.0227 0.8210 0.0395
-3.000 -0.1712 0.01819 0.00829 -0.0219 0.8086 0.0426
-2.750 -0.1465 0.01766 0.00760 -0.0210 0.7968 0.0478
-2.500 -0.1231 0.01697 0.00689 -0.0200 0.7858 0.0537
-2.250 -0.0990 0.01638 0.00621 -0.0190 0.7757 0.0637
-2.000 -0.0749 0.01576 0.00567 -0.0181 0.7654 0.0931
-1.750 -0.0517 0.01484 0.00527 -0.0174 0.7553 0.2044
-1.500 0.0298 0.01270 0.00549 -0.0263 0.7484 0.9345
-1.250 0.1242 0.01281 0.00519 -0.0394 0.7384 1.0000
-1.000 0.1474 0.01282 0.00501 -0.0386 0.7285 1.0000
-0.750 0.1706 0.01283 0.00482 -0.0377 0.7196 1.0000
-0.500 0.1948 0.01286 0.00471 -0.0371 0.7093 1.0000
-0.250 0.2188 0.01290 0.00461 -0.0364 0.7005 1.0000
0.000 0.2426 0.01294 0.00453 -0.0356 0.6918 1.0000
0.250 0.2667 0.01301 0.00449 -0.0350 0.6828 1.0000
0.500 0.2905 0.01308 0.00444 -0.0342 0.6752 1.0000
0.750 0.3147 0.01316 0.00447 -0.0336 0.6660 1.0000
1.000 0.3387 0.01325 0.00449 -0.0328 0.6585 1.0000
1.250 0.3629 0.01336 0.00454 -0.0322 0.6502 1.0000
1.500 0.3871 0.01347 0.00461 -0.0315 0.6425 1.0000
1.750 0.4113 0.01359 0.00469 -0.0308 0.6348 1.0000
2.000 0.4357 0.01372 0.00483 -0.0302 0.6271 1.0000
2.250 0.4599 0.01386 0.00495 -0.0295 0.6199 1.0000
2.500 0.4843 0.01401 0.00512 -0.0288 0.6122 1.0000
2.750 0.5086 0.01417 0.00528 -0.0281 0.6052 1.0000
3.000 0.5329 0.01435 0.00552 -0.0275 0.5976 1.0000
3.250 0.5572 0.01452 0.00571 -0.0268 0.5907 1.0000
3.500 0.5816 0.01472 0.00599 -0.0262 0.5831 1.0000
3.750 0.6059 0.01491 0.00622 -0.0255 0.5762 1.0000
4.000 0.6301 0.01513 0.00658 -0.0249 0.5685 1.0000
4.250 0.6544 0.01533 0.00685 -0.0242 0.5615 1.0000
4.500 0.6785 0.01558 0.00723 -0.0235 0.5532 1.0000
4.750 0.7024 0.01574 0.00746 -0.0226 0.5437 1.0000
5.000 0.7250 0.01577 0.00760 -0.0214 0.5247 1.0000
5.250 0.7470 0.01572 0.00755 -0.0200 0.5002 1.0000
5.500 0.7692 0.01579 0.00770 -0.0187 0.4757 1.0000
5.750 0.7918 0.01595 0.00798 -0.0177 0.4542 1.0000
6.000 0.8139 0.01614 0.00830 -0.0165 0.4291 1.0000
6.250 0.8345 0.01641 0.00858 -0.0151 0.3869 1.0000
6.500 0.8487 0.01724 0.00891 -0.0131 0.2785 1.0000
6.750 0.8526 0.01951 0.01015 -0.0106 0.1373 1.0000
7.000 0.8638 0.02105 0.01136 -0.0088 0.0853 1.0000
7.250 0.8764 0.02238 0.01258 -0.0069 0.0604 1.0000
7.500 0.8880 0.02373 0.01395 -0.0048 0.0461 1.0000
7.750 0.9000 0.02498 0.01525 -0.0028 0.0380 1.0000
8.000 0.9106 0.02627 0.01659 -0.0007 0.0334 1.0000
8.250 0.9192 0.02774 0.01813 0.0017 0.0311 1.0000
8.500 0.9298 0.02908 0.01961 0.0039 0.0293 1.0000
8.750 0.9410 0.03044 0.02108 0.0060 0.0273 1.0000
9.000 0.9520 0.03177 0.02246 0.0078 0.0251 1.0000
9.250 0.9618 0.03338 0.02414 0.0096 0.0236 1.0000
9.500 0.9752 0.03523 0.02609 0.0112 0.0225 1.0000
9.750 0.9907 0.03713 0.02819 0.0127 0.0217 1.0000
10.000 1.0056 0.03933 0.03063 0.0141 0.0210 1.0000
10.250 1.0179 0.04165 0.03320 0.0155 0.0206 1.0000
10.500 1.0257 0.04422 0.03605 0.0173 0.0201 1.0000
10.750 1.0288 0.04675 0.03885 0.0193 0.0197 1.0000
11.000 1.0280 0.04928 0.04163 0.0211 0.0192 1.0000
11.250 1.0252 0.05186 0.04443 0.0225 0.0187 1.0000
11.500 1.0204 0.05462 0.04740 0.0234 0.0183 1.0000
11.750 1.0137 0.05765 0.05064 0.0239 0.0179 1.0000
12.000 1.0048 0.06114 0.05438 0.0239 0.0177 1.0000
12.250 0.9961 0.06457 0.05796 0.0234 0.0173 1.0000
12.500 0.9835 0.06889 0.06248 0.0223 0.0172 1.0000
12.750 0.9703 0.07348 0.06725 0.0207 0.0171 1.0000
13.000 0.9519 0.07927 0.07326 0.0181 0.0172 1.0000
13.250 0.9266 0.08683 0.08105 0.0141 0.0176 1.0000
13.500 0.8915 0.09728 0.09176 0.0077 0.0183 1.0000
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Polar data table (+)
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