BOEING 707 .40 SPAN AIRFOIL (b707c-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: BOEING 707 .40 SPAN AIRFOIL (b707c-il) Reynolds number: 200,000 Max Cl/Cd: 43.24 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-b707c-il-200000.txt Download as CSV file: xf-b707c-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: BOEING 707 .40 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.3738 0.09788 0.09429 -0.0251 1.0000 0.0315
-10.000 -0.3731 0.09422 0.09064 -0.0260 1.0000 0.0320
-9.750 -0.4794 0.09871 0.09493 -0.0267 1.0000 0.0302
-9.500 -0.4752 0.09537 0.09160 -0.0267 1.0000 0.0306
-9.250 -0.4739 0.09185 0.08812 -0.0279 1.0000 0.0308
-9.000 -0.4733 0.08835 0.08465 -0.0290 1.0000 0.0312
-8.750 -0.4753 0.08462 0.08096 -0.0308 1.0000 0.0315
-8.500 -0.4820 0.08056 0.07694 -0.0336 1.0000 0.0317
-8.250 -0.4937 0.07704 0.07344 -0.0347 1.0000 0.0319
-8.000 -0.5010 0.07370 0.07009 -0.0348 1.0000 0.0322
-7.750 -0.5054 0.07049 0.06685 -0.0346 1.0000 0.0325
-7.500 -0.5081 0.06744 0.06375 -0.0339 1.0000 0.0333
-7.250 -0.5091 0.06449 0.06073 -0.0330 1.0000 0.0340
-7.000 -0.5082 0.06174 0.05787 -0.0316 1.0000 0.0348
-6.750 -0.5056 0.06115 0.05684 -0.0291 1.0000 0.0365
-6.500 -0.5038 0.05998 0.05533 -0.0261 1.0000 0.0367
-6.250 -0.4987 0.05543 0.05076 -0.0247 1.0000 0.0371
-6.000 -0.4920 0.05111 0.04654 -0.0235 1.0000 0.0376
-5.750 -0.4845 0.04819 0.04360 -0.0219 1.0000 0.0382
-5.500 -0.4759 0.04573 0.04108 -0.0201 1.0000 0.0390
-5.250 -0.4660 0.04353 0.03879 -0.0182 1.0000 0.0400
-5.000 -0.4549 0.04146 0.03658 -0.0163 1.0000 0.0413
-4.750 -0.4419 0.03974 0.03467 -0.0142 1.0000 0.0431
-4.500 -0.3998 0.03772 0.03202 -0.0167 0.9943 0.0462
-4.250 -0.3660 0.03379 0.02813 -0.0201 0.9885 0.0480
-4.000 -0.3284 0.03153 0.02569 -0.0229 0.9828 0.0514
-3.750 -0.2900 0.02942 0.02316 -0.0250 0.9756 0.0576
-3.500 -0.2528 0.02735 0.02100 -0.0276 0.9686 0.0638
-3.250 -0.2145 0.02532 0.01881 -0.0302 0.9609 0.0740
-3.000 -0.1787 0.02361 0.01696 -0.0326 0.9516 0.0961
-2.000 -0.0162 0.01701 0.00952 -0.0381 0.9120 0.0701
-1.750 0.0218 0.01569 0.00809 -0.0386 0.8972 0.0525
-1.500 0.0576 0.01467 0.00691 -0.0392 0.8793 0.0469
-1.250 0.0907 0.01391 0.00604 -0.0394 0.8462 0.0461
-1.000 0.1164 0.01292 0.00497 -0.0385 0.8005 0.0480
-0.750 0.1401 0.01241 0.00433 -0.0373 0.7625 0.0532
-0.500 0.1630 0.01200 0.00373 -0.0359 0.7356 0.0566
-0.250 0.1854 0.01170 0.00315 -0.0345 0.7033 0.0671
0.000 0.2053 0.01095 0.00284 -0.0329 0.6799 0.2617
0.250 0.2102 0.00943 0.00298 -0.0281 0.6602 0.7445
0.500 0.2308 0.00934 0.00312 -0.0253 0.6362 0.8579
0.750 0.2616 0.00953 0.00320 -0.0251 0.6002 0.9084
1.000 0.3074 0.00978 0.00333 -0.0282 0.5616 0.9474
1.250 0.3501 0.01003 0.00336 -0.0310 0.5047 0.9649
1.500 0.3977 0.01148 0.00358 -0.0357 0.2514 0.9859
1.750 0.4500 0.01221 0.00391 -0.0414 0.2001 0.9975
2.000 0.4733 0.01269 0.00399 -0.0412 0.0905 1.0000
2.250 0.4880 0.01297 0.00414 -0.0391 0.0690 1.0000
2.500 0.5045 0.01317 0.00427 -0.0372 0.0637 1.0000
2.750 0.5218 0.01340 0.00445 -0.0354 0.0596 1.0000
3.000 0.5405 0.01363 0.00468 -0.0337 0.0578 1.0000
3.250 0.5598 0.01390 0.00494 -0.0322 0.0563 1.0000
3.500 0.5799 0.01414 0.00520 -0.0308 0.0556 1.0000
3.750 0.6005 0.01441 0.00549 -0.0295 0.0550 1.0000
4.000 0.6212 0.01471 0.00583 -0.0283 0.0545 1.0000
4.250 0.6418 0.01503 0.00618 -0.0270 0.0532 1.0000
4.500 0.6623 0.01538 0.00655 -0.0257 0.0517 1.0000
4.750 0.6824 0.01578 0.00697 -0.0244 0.0503 1.0000
5.000 0.7014 0.01631 0.00752 -0.0229 0.0481 1.0000
5.250 0.7180 0.01706 0.00826 -0.0211 0.0461 1.0000
5.500 0.7383 0.01753 0.00878 -0.0198 0.0453 1.0000
5.750 0.7578 0.01812 0.00942 -0.0184 0.0443 1.0000
6.000 0.7773 0.01874 0.01008 -0.0171 0.0418 1.0000
6.250 0.7946 0.01970 0.01102 -0.0155 0.0390 1.0000
6.500 0.8125 0.02104 0.01237 -0.0139 0.0364 1.0000
6.750 0.8332 0.02218 0.01357 -0.0127 0.0336 1.0000
7.000 0.8549 0.02350 0.01486 -0.0119 0.0321 1.0000
7.250 0.8804 0.02596 0.01726 -0.0119 0.0308 1.0000
7.500 0.9025 0.02729 0.01882 -0.0109 0.0303 1.0000
7.750 0.9249 0.02911 0.02086 -0.0100 0.0300 1.0000
8.000 0.9460 0.03123 0.02321 -0.0089 0.0299 1.0000
8.250 0.9657 0.03373 0.02597 -0.0078 0.0301 1.0000
8.500 0.9844 0.03693 0.02935 -0.0069 0.0305 1.0000
8.750 1.0024 0.03871 0.03141 -0.0053 0.0315 1.0000
11.750 0.6529 0.10823 0.10474 -0.0023 0.0649 1.0000
12.000 0.6382 0.11499 0.11147 -0.0062 0.0625 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING 707 .40 SPAN AIRFOIL (b707c-il)