B-29 TIP AIRFOIL (b29tip-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: B-29 TIP AIRFOIL (b29tip-il) Reynolds number: 1,000,000 Max Cl/Cd: 106.14 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-b29tip-il-1000000.txt Download as CSV file: xf-b29tip-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5448 0.08199 0.08049 -0.0099 1.0000 0.0134
-8.750 -0.5491 0.07788 0.07640 -0.0128 1.0000 0.0135
-8.500 -0.5594 0.07369 0.07223 -0.0159 1.0000 0.0136
-8.250 -0.5635 0.07001 0.06853 -0.0175 1.0000 0.0137
-8.000 -0.5638 0.06561 0.06410 -0.0196 1.0000 0.0138
-7.750 -0.5600 0.06181 0.06024 -0.0211 1.0000 0.0141
-7.500 -0.5547 0.05781 0.05618 -0.0222 1.0000 0.0144
-7.250 -0.5324 0.05191 0.04999 -0.0275 0.9079 0.0149
-7.000 -0.5270 0.04798 0.04578 -0.0267 0.8675 0.0155
-6.750 -0.5184 0.04304 0.04054 -0.0257 0.8431 0.0166
-5.250 -0.4410 0.01864 0.01377 -0.0144 0.7587 0.0166
-5.000 -0.4199 0.01625 0.01102 -0.0130 0.7484 0.0161
-4.750 -0.3964 0.01464 0.00912 -0.0120 0.7382 0.0159
-4.500 -0.3715 0.01345 0.00774 -0.0113 0.7281 0.0158
-4.250 -0.3459 0.01270 0.00686 -0.0108 0.7186 0.0162
-4.000 -0.3206 0.01186 0.00588 -0.0101 0.7094 0.0163
-3.750 -0.2951 0.01115 0.00506 -0.0095 0.6999 0.0165
-3.500 -0.2694 0.01064 0.00446 -0.0090 0.6910 0.0169
-3.250 -0.2433 0.01027 0.00401 -0.0085 0.6821 0.0173
-3.000 -0.2169 0.00998 0.00367 -0.0081 0.6732 0.0176
-2.750 -0.1929 0.00927 0.00283 -0.0073 0.6649 0.0184
-2.500 -0.1669 0.00892 0.00242 -0.0068 0.6562 0.0192
-2.250 -0.1405 0.00868 0.00213 -0.0064 0.6480 0.0201
-2.000 -0.1138 0.00850 0.00189 -0.0061 0.6395 0.0216
-1.750 -0.0868 0.00835 0.00171 -0.0059 0.6313 0.0234
-1.250 -0.0343 0.00784 0.00131 -0.0051 0.6154 0.0697
-1.000 -0.0083 0.00764 0.00121 -0.0047 0.6078 0.1116
-0.750 0.0169 0.00728 0.00112 -0.0043 0.6002 0.1991
-0.500 0.0399 0.00676 0.00105 -0.0036 0.5932 0.3481
-0.250 0.0544 0.00562 0.00093 -0.0011 0.5865 0.6629
0.000 0.0709 0.00487 0.00099 0.0017 0.5801 0.9056
0.250 0.1151 0.00498 0.00114 -0.0015 0.5730 0.9525
0.500 0.1546 0.00518 0.00128 -0.0038 0.5661 0.9683
0.750 0.1958 0.00531 0.00136 -0.0066 0.5590 0.9729
1.000 0.2483 0.00559 0.00156 -0.0118 0.5519 0.9810
1.250 0.2989 0.00575 0.00169 -0.0167 0.5451 0.9860
1.500 0.3378 0.00582 0.00170 -0.0191 0.5383 0.9884
1.750 0.3769 0.00587 0.00173 -0.0216 0.5321 0.9916
2.000 0.4118 0.00594 0.00177 -0.0231 0.5256 0.9943
2.250 0.4457 0.00593 0.00175 -0.0245 0.5196 0.9955
2.500 0.4783 0.00592 0.00172 -0.0257 0.5122 0.9967
2.750 0.5106 0.00593 0.00171 -0.0267 0.5011 0.9980
3.000 0.5427 0.00596 0.00171 -0.0278 0.4901 0.9992
3.250 0.5729 0.00600 0.00170 -0.0284 0.4738 1.0000
3.500 0.5975 0.00606 0.00174 -0.0278 0.4620 1.0000
3.750 0.6222 0.00613 0.00179 -0.0272 0.4512 1.0000
4.000 0.6469 0.00621 0.00185 -0.0267 0.4351 1.0000
4.250 0.6713 0.00634 0.00191 -0.0261 0.4105 1.0000
4.500 0.6952 0.00655 0.00200 -0.0254 0.3790 1.0000
4.750 0.7188 0.00681 0.00214 -0.0248 0.3412 1.0000
5.000 0.7407 0.00737 0.00239 -0.0240 0.2696 1.0000
5.250 0.7597 0.00836 0.00288 -0.0230 0.1612 1.0000
5.500 0.7802 0.00908 0.00330 -0.0220 0.1008 1.0000
5.750 0.8017 0.00960 0.00364 -0.0211 0.0651 1.0000
6.000 0.8231 0.01010 0.00402 -0.0202 0.0389 1.0000
6.250 0.8436 0.01072 0.00450 -0.0190 0.0178 1.0000
6.500 0.8657 0.01110 0.00492 -0.0180 0.0150 1.0000
6.750 0.8874 0.01150 0.00533 -0.0170 0.0132 1.0000
7.000 0.9071 0.01214 0.00606 -0.0157 0.0116 1.0000
7.250 0.9285 0.01254 0.00650 -0.0146 0.0113 1.0000
7.500 0.9491 0.01299 0.00701 -0.0135 0.0108 1.0000
7.750 0.9691 0.01350 0.00756 -0.0122 0.0102 1.0000
8.000 0.9884 0.01404 0.00816 -0.0109 0.0098 1.0000
8.250 1.0070 0.01463 0.00881 -0.0095 0.0094 1.0000
8.500 1.0240 0.01536 0.00958 -0.0078 0.0089 1.0000
8.750 1.0387 0.01625 0.01055 -0.0059 0.0087 1.0000
9.000 1.0448 0.01792 0.01235 -0.0028 0.0082 1.0000
9.250 1.0642 0.01831 0.01279 -0.0016 0.0079 1.0000
9.500 1.0792 0.01911 0.01366 0.0003 0.0078 1.0000
9.750 1.0930 0.01999 0.01463 0.0022 0.0076 1.0000
10.000 1.1058 0.02094 0.01566 0.0042 0.0073 1.0000
10.250 1.1175 0.02197 0.01678 0.0063 0.0072 1.0000
10.500 1.1284 0.02305 0.01795 0.0084 0.0071 1.0000
10.750 1.1366 0.02414 0.01913 0.0110 0.0069 1.0000
11.000 1.1415 0.02531 0.02039 0.0139 0.0069 1.0000
11.250 1.1462 0.02672 0.02192 0.0164 0.0068 1.0000
11.500 1.1521 0.02808 0.02338 0.0182 0.0066 1.0000
11.750 1.1561 0.02994 0.02538 0.0200 0.0066 1.0000
12.000 1.1637 0.03101 0.02652 0.0209 0.0063 1.0000
12.250 1.1653 0.03329 0.02897 0.0223 0.0064 1.0000
12.500 1.1651 0.03583 0.03167 0.0234 0.0064 1.0000
12.750 1.1718 0.03714 0.03299 0.0234 0.0060 1.0000
13.000 1.1654 0.04061 0.03670 0.0242 0.0062 1.0000
13.250 1.1627 0.04335 0.03948 0.0241 0.0058 1.0000
13.500 1.1515 0.04745 0.04377 0.0239 0.0058 1.0000
13.750 1.1525 0.05003 0.04647 0.0234 0.0059 1.0000
14.000 1.1276 0.05623 0.05291 0.0221 0.0057 1.0000
14.250 1.1045 0.06246 0.05936 0.0202 0.0057 1.0000
14.500 1.0873 0.06815 0.06522 0.0178 0.0057 1.0000
14.750 1.1000 0.06947 0.06657 0.0169 0.0058 1.0000
15.000 1.0672 0.07830 0.07561 0.0128 0.0057 1.0000
15.250 1.0444 0.08610 0.08357 0.0088 0.0057 1.0000
15.500 1.0387 0.09127 0.08884 0.0059 0.0058 1.0000
15.750 1.0245 0.09833 0.09602 0.0019 0.0058 1.0000
16.000 0.9991 0.10813 0.10597 -0.0036 0.0058 1.0000
16.250 0.9824 0.11675 0.11471 -0.0084 0.0059 1.0000
|
Polar data table (+)
Polar graphs
<< Back to B-29 TIP AIRFOIL (b29tip-il)