B-29 TIP AIRFOIL (b29tip-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: B-29 TIP AIRFOIL (b29tip-il) Reynolds number: 50,000 Max Cl/Cd: 34.8 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-b29tip-il-50000-n5.txt Download as CSV file: xf-b29tip-il-50000-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 = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.4985 0.09455 0.08808 -0.0078 1.0000 0.1241
-8.000 -0.5008 0.09068 0.08429 -0.0105 1.0000 0.1233
-7.750 -0.5125 0.07992 0.07337 -0.0212 1.0000 0.0598
-7.500 -0.5072 0.07596 0.06941 -0.0219 1.0000 0.0583
-7.250 -0.5035 0.07181 0.06522 -0.0233 1.0000 0.0574
-7.000 -0.4990 0.06765 0.06098 -0.0245 1.0000 0.0568
-6.750 -0.4931 0.06356 0.05678 -0.0255 1.0000 0.0564
-6.500 -0.4859 0.05958 0.05263 -0.0260 1.0000 0.0566
-6.250 -0.4769 0.05577 0.04862 -0.0262 1.0000 0.0564
-6.000 -0.4667 0.05211 0.04472 -0.0259 1.0000 0.0560
-5.750 -0.4555 0.04858 0.04093 -0.0252 1.0000 0.0552
-5.500 -0.4439 0.04528 0.03728 -0.0242 1.0000 0.0545
-5.250 -0.4325 0.04230 0.03398 -0.0226 1.0000 0.0539
-5.000 -0.4223 0.03978 0.03115 -0.0206 1.0000 0.0536
-4.750 -0.4139 0.03760 0.02867 -0.0181 1.0000 0.0535
-4.500 -0.4003 0.03546 0.02617 -0.0165 0.9958 0.0536
-4.250 -0.3638 0.03268 0.02284 -0.0187 0.9794 0.0543
-4.000 -0.3251 0.03038 0.02003 -0.0210 0.9657 0.0568
-3.750 -0.2848 0.02855 0.01756 -0.0232 0.9528 0.0609
-3.500 -0.2459 0.02664 0.01552 -0.0255 0.9404 0.0645
-3.250 -0.2051 0.02509 0.01376 -0.0277 0.9286 0.0689
-3.000 -0.1631 0.02376 0.01217 -0.0299 0.9174 0.0767
-2.750 -0.1243 0.02267 0.01097 -0.0319 0.9055 0.0928
-2.500 -0.0911 0.02159 0.00991 -0.0329 0.8923 0.1211
-2.250 -0.0646 0.02018 0.00908 -0.0330 0.8790 0.2087
-2.000 0.0765 0.01731 0.00824 -0.0518 0.8858 1.0000
-1.750 0.1035 0.01737 0.00793 -0.0518 0.8688 1.0000
-1.500 0.1283 0.01747 0.00773 -0.0513 0.8532 1.0000
-1.250 0.1519 0.01760 0.00760 -0.0505 0.8385 1.0000
-1.000 0.1747 0.01775 0.00752 -0.0497 0.8248 1.0000
-0.750 0.1973 0.01791 0.00746 -0.0487 0.8117 1.0000
-0.500 0.2198 0.01809 0.00746 -0.0478 0.7994 1.0000
-0.250 0.2423 0.01828 0.00749 -0.0468 0.7879 1.0000
0.000 0.2649 0.01847 0.00752 -0.0457 0.7775 1.0000
0.250 0.2873 0.01870 0.00765 -0.0448 0.7660 1.0000
0.500 0.3098 0.01894 0.00779 -0.0440 0.7553 1.0000
0.750 0.3325 0.01917 0.00793 -0.0430 0.7458 1.0000
1.000 0.3552 0.01942 0.00813 -0.0420 0.7360 1.0000
1.250 0.3777 0.01972 0.00839 -0.0412 0.7258 1.0000
1.500 0.4005 0.01998 0.00860 -0.0403 0.7171 1.0000
1.750 0.4232 0.02029 0.00889 -0.0394 0.7076 1.0000
2.000 0.4463 0.02063 0.00927 -0.0387 0.6982 1.0000
2.250 0.4697 0.02089 0.00951 -0.0377 0.6905 1.0000
2.500 0.4921 0.02131 0.00999 -0.0371 0.6806 1.0000
2.750 0.5149 0.02166 0.01038 -0.0362 0.6722 1.0000
3.000 0.5377 0.02203 0.01083 -0.0353 0.6637 1.0000
3.250 0.5599 0.02249 0.01138 -0.0346 0.6546 1.0000
3.500 0.5831 0.02280 0.01175 -0.0335 0.6474 1.0000
3.750 0.6046 0.02335 0.01244 -0.0328 0.6377 1.0000
4.000 0.6274 0.02375 0.01298 -0.0318 0.6299 1.0000
4.250 0.6493 0.02425 0.01363 -0.0310 0.6209 1.0000
4.500 0.6708 0.02480 0.01434 -0.0301 0.6118 1.0000
4.750 0.6941 0.02514 0.01483 -0.0290 0.6042 1.0000
5.000 0.7143 0.02579 0.01573 -0.0280 0.5937 1.0000
5.250 0.7359 0.02627 0.01641 -0.0269 0.5838 1.0000
5.500 0.7580 0.02643 0.01676 -0.0253 0.5716 1.0000
5.750 0.7785 0.02631 0.01681 -0.0231 0.5524 1.0000
6.000 0.8000 0.02565 0.01626 -0.0202 0.5279 1.0000
6.250 0.8194 0.02514 0.01593 -0.0175 0.4987 1.0000
6.500 0.8372 0.02491 0.01588 -0.0149 0.4660 1.0000
6.750 0.8542 0.02482 0.01596 -0.0124 0.4263 1.0000
7.000 0.8672 0.02492 0.01604 -0.0094 0.3560 1.0000
7.250 0.8684 0.02632 0.01647 -0.0055 0.2192 1.0000
7.500 0.8635 0.02906 0.01841 -0.0026 0.1484 1.0000
7.750 0.8623 0.03149 0.02051 0.0000 0.1128 1.0000
8.000 0.8619 0.03373 0.02259 0.0025 0.0925 1.0000
8.250 0.8637 0.03576 0.02458 0.0048 0.0780 1.0000
8.500 0.8656 0.03766 0.02645 0.0072 0.0687 1.0000
8.750 0.8711 0.03945 0.02835 0.0095 0.0609 1.0000
9.000 0.8790 0.04133 0.03019 0.0115 0.0561 1.0000
9.250 0.9017 0.04299 0.03210 0.0132 0.0510 1.0000
9.500 0.9201 0.04476 0.03402 0.0145 0.0461 1.0000
9.750 0.9442 0.04707 0.03645 0.0153 0.0424 1.0000
10.000 0.9680 0.04995 0.03978 0.0161 0.0405 1.0000
10.250 0.9808 0.05297 0.04317 0.0173 0.0394 1.0000
10.500 0.9851 0.05615 0.04671 0.0189 0.0386 1.0000
10.750 0.9832 0.05929 0.05015 0.0204 0.0379 1.0000
11.000 0.9768 0.06259 0.05372 0.0216 0.0371 1.0000
11.250 0.9676 0.06619 0.05757 0.0220 0.0367 1.0000
11.500 0.9557 0.07012 0.06173 0.0219 0.0363 1.0000
11.750 0.9410 0.07461 0.06644 0.0209 0.0362 1.0000
12.000 0.9233 0.07978 0.07182 0.0190 0.0363 1.0000
12.250 0.9014 0.08609 0.07833 0.0160 0.0370 1.0000
12.500 0.8784 0.09325 0.08564 0.0119 0.0376 1.0000
|
Polar data table (+)
Polar graphs
<< Back to B-29 TIP AIRFOIL (b29tip-il)