BOEING 707 .08 SPAN AIRFOIL (b707a-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: BOEING 707 .08 SPAN AIRFOIL (b707a-il) Reynolds number: 50,000 Max Cl/Cd: 20.68 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-b707a-il-50000-n5.txt Download as CSV file: xf-b707a-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: BOEING 707 .08 SPAN 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
------ -------- --------- --------- -------- -------- --------
-12.250 -0.7901 0.09362 0.08526 -0.0148 1.0000 0.0499
-12.000 -0.8178 0.08588 0.07737 -0.0195 1.0000 0.0495
-11.750 -0.8432 0.07937 0.07069 -0.0231 1.0000 0.0492
-11.500 -0.8666 0.07388 0.06500 -0.0255 1.0000 0.0491
-11.250 -0.8855 0.06945 0.06037 -0.0265 1.0000 0.0490
-11.000 -0.9021 0.06565 0.05634 -0.0262 1.0000 0.0490
-10.750 -0.9151 0.06241 0.05288 -0.0248 1.0000 0.0492
-10.500 -0.9251 0.05943 0.04965 -0.0227 1.0000 0.0495
-10.250 -0.9290 0.05645 0.04637 -0.0208 1.0000 0.0500
-10.000 -0.9286 0.05351 0.04309 -0.0189 1.0000 0.0507
-9.750 -0.9232 0.05071 0.03994 -0.0172 1.0000 0.0517
-9.500 -0.9073 0.04852 0.03768 -0.0162 1.0000 0.0530
-9.250 -0.8922 0.04635 0.03534 -0.0150 1.0000 0.0550
-9.000 -0.8784 0.04410 0.03268 -0.0134 1.0000 0.0579
-8.750 -0.8571 0.04234 0.03100 -0.0127 1.0000 0.0613
-8.500 -0.8360 0.04048 0.02894 -0.0115 1.0000 0.0660
-8.250 -0.8130 0.03882 0.02723 -0.0105 1.0000 0.0721
-8.000 -0.7916 0.03726 0.02575 -0.0092 1.0000 0.0784
-7.750 -0.7759 0.03586 0.02434 -0.0075 1.0000 0.0856
-7.500 -0.7650 0.03451 0.02303 -0.0056 1.0000 0.0952
-7.250 -0.7548 0.03320 0.02178 -0.0036 1.0000 0.1058
-7.000 -0.7431 0.03198 0.02057 -0.0016 1.0000 0.1159
-6.750 -0.7316 0.03076 0.01941 0.0005 1.0000 0.1245
-6.500 -0.7192 0.02961 0.01828 0.0026 1.0000 0.1310
-6.250 -0.7059 0.02852 0.01718 0.0047 1.0000 0.1364
-6.000 -0.6922 0.02745 0.01610 0.0067 1.0000 0.1430
-5.750 -0.6787 0.02631 0.01501 0.0086 1.0000 0.1522
-5.500 -0.6645 0.02513 0.01394 0.0104 1.0000 0.1657
-5.250 -0.6505 0.02379 0.01282 0.0122 1.0000 0.1944
-5.000 -0.6383 0.02241 0.01197 0.0143 1.0000 0.2610
-4.750 -0.6243 0.02231 0.01232 0.0172 1.0000 0.3993
-4.500 -0.6060 0.02228 0.01216 0.0194 1.0000 0.4389
-4.250 -0.5876 0.02232 0.01218 0.0217 1.0000 0.4690
-4.000 -0.5673 0.02225 0.01206 0.0236 1.0000 0.4897
-3.750 -0.5463 0.02207 0.01174 0.0251 1.0000 0.5053
-3.500 -0.5239 0.02195 0.01165 0.0268 1.0000 0.5188
-3.250 -0.5018 0.02175 0.01139 0.0282 1.0000 0.5312
-3.000 -0.4796 0.02155 0.01114 0.0295 1.0000 0.5435
-2.750 -0.4567 0.02134 0.01096 0.0309 1.0000 0.5540
-2.500 -0.4346 0.02112 0.01069 0.0321 1.0000 0.5662
-2.250 -0.4121 0.02098 0.01064 0.0337 1.0000 0.5790
-2.000 -0.3904 0.02081 0.01053 0.0352 1.0000 0.5933
-1.750 -0.3683 0.02064 0.01040 0.0366 1.0000 0.6066
-1.500 -0.3449 0.02043 0.01026 0.0377 1.0000 0.6158
-1.250 -0.3213 0.02017 0.01003 0.0385 1.0000 0.6222
-1.000 -0.2970 0.01989 0.00974 0.0390 1.0000 0.6270
-0.750 -0.2724 0.01963 0.00950 0.0393 1.0000 0.6311
-0.500 -0.2476 0.01941 0.00937 0.0398 1.0000 0.6346
-0.250 -0.2020 0.01927 0.00933 0.0360 0.9776 0.6390
0.000 -0.1544 0.01900 0.00916 0.0323 0.9159 0.6445
0.250 -0.0590 0.01901 0.00876 0.0207 0.7026 0.6516
0.500 -0.0235 0.01944 0.00865 0.0201 0.5895 0.6575
0.750 0.0025 0.01976 0.00864 0.0209 0.5339 0.6651
1.000 0.0285 0.01998 0.00868 0.0214 0.4936 0.6740
1.250 0.0547 0.02014 0.00877 0.0220 0.4597 0.6839
1.500 0.0808 0.02027 0.00885 0.0226 0.4299 0.6961
1.750 0.1066 0.02044 0.00893 0.0232 0.4053 0.7105
2.000 0.1330 0.02065 0.00911 0.0238 0.3837 0.7273
2.250 0.1606 0.02096 0.00940 0.0241 0.3643 0.7483
2.500 0.1912 0.02138 0.00985 0.0239 0.3457 0.7747
2.750 0.2271 0.02191 0.01049 0.0228 0.3275 0.8063
3.000 0.2686 0.02257 0.01128 0.0206 0.3097 0.8439
3.250 0.3148 0.02332 0.01213 0.0173 0.2922 0.8852
3.500 0.3604 0.02397 0.01299 0.0140 0.2748 0.9282
3.750 0.4043 0.02441 0.01363 0.0107 0.2551 0.9685
4.000 0.4451 0.02441 0.01371 0.0078 0.2346 1.0000
4.250 0.4629 0.02436 0.01353 0.0094 0.2215 1.0000
4.500 0.4817 0.02449 0.01347 0.0108 0.2098 1.0000
4.750 0.5023 0.02499 0.01389 0.0120 0.1988 1.0000
5.000 0.5233 0.02572 0.01466 0.0132 0.1879 1.0000
5.250 0.5441 0.02650 0.01537 0.0143 0.1783 1.0000
5.500 0.5649 0.02734 0.01624 0.0155 0.1687 1.0000
5.750 0.5857 0.02839 0.01740 0.0166 0.1599 1.0000
6.000 0.6065 0.02933 0.01834 0.0177 0.1535 1.0000
6.250 0.6267 0.03066 0.01986 0.0189 0.1472 1.0000
6.500 0.6468 0.03182 0.02111 0.0200 0.1419 1.0000
6.750 0.6673 0.03289 0.02215 0.0210 0.1380 1.0000
7.000 0.6846 0.03472 0.02436 0.0223 0.1335 1.0000
7.250 0.7024 0.03630 0.02614 0.0234 0.1301 1.0000
7.500 0.7207 0.03779 0.02773 0.0245 0.1278 1.0000
7.750 0.7386 0.03933 0.02935 0.0256 0.1256 1.0000
8.000 0.7469 0.04224 0.03277 0.0270 0.1227 1.0000
8.250 0.7562 0.04473 0.03557 0.0282 0.1196 1.0000
8.500 0.7730 0.04581 0.03670 0.0293 0.1161 1.0000
8.750 0.7832 0.04780 0.03886 0.0304 0.1129 1.0000
9.000 0.7755 0.05219 0.04371 0.0315 0.1107 1.0000
9.250 0.7626 0.05700 0.04884 0.0321 0.1093 1.0000
9.500 0.7319 0.06387 0.05596 0.0313 0.1095 1.0000
9.750 0.6869 0.07267 0.06482 0.0283 0.1115 1.0000
10.000 0.6530 0.08309 0.07520 0.0209 0.1129 1.0000
10.250 0.6325 0.09105 0.08311 0.0158 0.1127 1.0000
10.500 0.6174 0.09783 0.08983 0.0117 0.1119 1.0000
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Polar data table (+)
Polar graphs
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