BOEING 707 .54 SPAN AIRFOIL (b707d-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: BOEING 707 .54 SPAN AIRFOIL (b707d-il) Reynolds number: 200,000 Max Cl/Cd: 52.25 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-b707d-il-200000-n5.txt Download as CSV file: xf-b707d-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: BOEING 707 .54 SPAN 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4636 0.09244 0.08895 -0.0329 1.0000 0.0256
-9.000 -0.4668 0.08791 0.08448 -0.0370 1.0000 0.0256
-7.250 -0.4951 0.06003 0.05626 -0.0380 1.0000 0.0268
-7.000 -0.4945 0.05726 0.05353 -0.0362 1.0000 0.0277
-6.750 -0.4775 0.05371 0.04989 -0.0379 0.9943 0.0287
-6.500 -0.4542 0.04993 0.04591 -0.0407 0.9866 0.0299
-6.250 -0.4291 0.04668 0.04244 -0.0430 0.9789 0.0328
-6.000 -0.3944 0.04539 0.04063 -0.0444 0.9714 0.0371
-5.750 -0.3656 0.04313 0.03802 -0.0458 0.9645 0.0375
-5.500 -0.3392 0.04047 0.03505 -0.0468 0.9565 0.0376
-5.250 -0.3131 0.03788 0.03215 -0.0475 0.9484 0.0377
-5.000 -0.2931 0.03183 0.02583 -0.0487 0.9407 0.0392
-4.750 -0.2668 0.02930 0.02316 -0.0495 0.9330 0.0403
-4.500 -0.2384 0.02719 0.02084 -0.0496 0.9251 0.0340
-4.250 -0.2096 0.02539 0.01875 -0.0499 0.9181 0.0341
-4.000 -0.1833 0.02329 0.01634 -0.0495 0.9092 0.0308
-3.750 -0.1517 0.02235 0.01490 -0.0488 0.9021 0.0264
-3.500 -0.1249 0.02111 0.01356 -0.0487 0.8937 0.0252
-3.250 -0.0981 0.01949 0.01177 -0.0485 0.8863 0.0245
-3.000 -0.0715 0.01828 0.01040 -0.0482 0.8783 0.0240
-2.750 -0.0447 0.01725 0.00921 -0.0479 0.8707 0.0235
-2.500 -0.0178 0.01635 0.00820 -0.0476 0.8635 0.0230
-2.250 0.0084 0.01556 0.00733 -0.0471 0.8561 0.0226
-2.000 0.0346 0.01485 0.00656 -0.0467 0.8490 0.0222
-1.750 0.0600 0.01423 0.00590 -0.0461 0.8414 0.0219
-1.500 0.0853 0.01367 0.00529 -0.0456 0.8349 0.0216
-1.250 0.1100 0.01318 0.00479 -0.0449 0.8276 0.0214
-1.000 0.1350 0.01276 0.00433 -0.0443 0.8210 0.0213
-0.750 0.1598 0.01240 0.00396 -0.0438 0.8140 0.0212
-0.500 0.1852 0.01212 0.00362 -0.0433 0.8075 0.0212
-0.250 0.2109 0.01191 0.00334 -0.0429 0.8009 0.0213
0.000 0.2370 0.01175 0.00312 -0.0426 0.7941 0.0216
0.250 0.2634 0.01164 0.00293 -0.0422 0.7880 0.0222
0.500 0.2895 0.01155 0.00279 -0.0419 0.7801 0.0232
0.750 0.3153 0.01147 0.00264 -0.0413 0.7678 0.0251
1.000 0.3391 0.01133 0.00238 -0.0400 0.7402 0.0346
1.250 0.3516 0.00978 0.00218 -0.0373 0.7061 0.5890
1.500 0.3679 0.00920 0.00218 -0.0343 0.6701 0.7705
1.750 0.4161 0.00903 0.00229 -0.0378 0.6342 0.9356
2.000 0.4471 0.00915 0.00227 -0.0385 0.6039 0.9488
2.250 0.4783 0.00932 0.00228 -0.0393 0.5640 0.9636
2.500 0.5066 0.00970 0.00230 -0.0396 0.4856 0.9767
2.750 0.5332 0.01043 0.00251 -0.0399 0.3775 0.9893
3.000 0.5613 0.01104 0.00279 -0.0406 0.3078 1.0000
3.250 0.5816 0.01145 0.00302 -0.0394 0.2705 1.0000
3.500 0.6027 0.01181 0.00327 -0.0383 0.2422 1.0000
3.750 0.6244 0.01211 0.00353 -0.0373 0.2212 1.0000
4.000 0.6465 0.01240 0.00382 -0.0364 0.2046 1.0000
4.250 0.6678 0.01278 0.00413 -0.0353 0.1792 1.0000
4.500 0.6857 0.01348 0.00453 -0.0339 0.1217 1.0000
4.750 0.7019 0.01441 0.00509 -0.0322 0.0575 1.0000
5.000 0.7236 0.01482 0.00558 -0.0312 0.0483 1.0000
5.250 0.7458 0.01519 0.00603 -0.0303 0.0382 1.0000
5.500 0.7672 0.01567 0.00652 -0.0292 0.0277 1.0000
5.750 0.7877 0.01626 0.00711 -0.0280 0.0227 1.0000
6.000 0.8078 0.01693 0.00792 -0.0268 0.0200 1.0000
6.250 0.8262 0.01780 0.00894 -0.0252 0.0176 1.0000
6.500 0.8412 0.01898 0.01028 -0.0232 0.0150 1.0000
6.750 0.8603 0.01970 0.01112 -0.0219 0.0134 1.0000
7.000 0.8771 0.02068 0.01221 -0.0202 0.0125 1.0000
7.250 0.8932 0.02180 0.01342 -0.0185 0.0115 1.0000
7.500 0.9095 0.02298 0.01467 -0.0169 0.0110 1.0000
7.750 0.9255 0.02434 0.01613 -0.0153 0.0105 1.0000
8.000 0.9415 0.02607 0.01790 -0.0139 0.0101 1.0000
8.250 0.9598 0.02796 0.01992 -0.0127 0.0098 1.0000
8.500 0.9785 0.02972 0.02190 -0.0115 0.0096 1.0000
8.750 0.9958 0.03150 0.02401 -0.0100 0.0092 1.0000
9.000 1.0109 0.03325 0.02608 -0.0084 0.0085 1.0000
9.250 1.0239 0.03462 0.02767 -0.0067 0.0076 1.0000
9.500 1.0346 0.03631 0.02957 -0.0049 0.0073 1.0000
9.750 1.0417 0.03842 0.03194 -0.0028 0.0071 1.0000
10.000 1.0451 0.04047 0.03422 -0.0003 0.0069 1.0000
10.250 1.0434 0.04269 0.03665 0.0027 0.0068 1.0000
10.500 1.0385 0.04521 0.03941 0.0056 0.0067 1.0000
10.750 1.0306 0.04797 0.04240 0.0081 0.0067 1.0000
11.000 1.0198 0.05109 0.04577 0.0102 0.0066 1.0000
11.250 1.0064 0.05463 0.04953 0.0115 0.0066 1.0000
11.500 0.9936 0.05828 0.05341 0.0120 0.0066 1.0000
11.750 0.9776 0.06267 0.05801 0.0117 0.0066 1.0000
12.000 0.9615 0.06753 0.06308 0.0103 0.0066 1.0000
12.250 0.9431 0.07326 0.06899 0.0078 0.0066 1.0000
12.500 0.9244 0.07997 0.07591 0.0040 0.0067 1.0000
12.750 0.9033 0.08836 0.08450 -0.0015 0.0068 1.0000
13.000 0.8595 0.10602 0.10242 -0.0132 0.0072 1.0000
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Polar data table (+)
Polar graphs
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