BOEING-VERTOL VR-1 AIRFOIL (vr1-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: BOEING-VERTOL VR-1 AIRFOIL (vr1-il) Reynolds number: 100,000 Max Cl/Cd: 44.06 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-vr1-il-100000-n5.txt Download as CSV file: xf-vr1-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: BOEING-VERTOL VR-1 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.5818 0.07337 0.06761 -0.0543 1.0000 0.0350
-10.250 -0.5966 0.06849 0.06268 -0.0569 1.0000 0.0349
-10.000 -0.6134 0.06449 0.05863 -0.0578 1.0000 0.0347
-9.750 -0.6334 0.06115 0.05521 -0.0564 1.0000 0.0345
-9.500 -0.6503 0.05818 0.05214 -0.0540 1.0000 0.0343
-9.250 -0.6629 0.05513 0.04895 -0.0515 1.0000 0.0342
-9.000 -0.6738 0.05209 0.04571 -0.0486 1.0000 0.0340
-8.750 -0.6812 0.04921 0.04261 -0.0456 1.0000 0.0338
-8.500 -0.6864 0.04650 0.03965 -0.0423 1.0000 0.0337
-8.250 -0.6896 0.04396 0.03684 -0.0389 1.0000 0.0337
-8.000 -0.6911 0.04160 0.03418 -0.0354 1.0000 0.0339
-7.750 -0.6900 0.03948 0.03172 -0.0320 1.0000 0.0343
-7.500 -0.6864 0.03751 0.02941 -0.0288 1.0000 0.0346
-7.250 -0.6805 0.03567 0.02720 -0.0257 1.0000 0.0349
-7.000 -0.6549 0.03346 0.02454 -0.0263 0.9954 0.0351
-6.750 -0.6264 0.03136 0.02219 -0.0273 0.9903 0.0353
-6.500 -0.5962 0.02953 0.02015 -0.0285 0.9855 0.0356
-6.250 -0.5661 0.02798 0.01843 -0.0294 0.9806 0.0360
-6.000 -0.5358 0.02660 0.01690 -0.0303 0.9754 0.0365
-5.750 -0.5027 0.02535 0.01554 -0.0317 0.9713 0.0371
-5.500 -0.4738 0.02433 0.01445 -0.0322 0.9655 0.0383
-5.250 -0.4426 0.02340 0.01341 -0.0331 0.9602 0.0399
-5.000 -0.4087 0.02253 0.01243 -0.0345 0.9564 0.0412
-4.750 -0.3821 0.02162 0.01155 -0.0346 0.9499 0.0423
-4.500 -0.3521 0.02084 0.01076 -0.0354 0.9445 0.0436
-4.250 -0.3183 0.02014 0.01002 -0.0368 0.9407 0.0453
-4.000 -0.2924 0.01961 0.00941 -0.0366 0.9335 0.0475
-3.750 -0.2618 0.01903 0.00879 -0.0373 0.9282 0.0513
-3.500 -0.2279 0.01849 0.00822 -0.0386 0.9245 0.0582
-3.000 -0.1772 0.01702 0.00730 -0.0381 0.9110 0.1526
-2.750 -0.1542 0.01579 0.00702 -0.0379 0.9056 0.3463
-2.500 -0.1385 0.01485 0.00709 -0.0355 0.8977 0.5493
-2.250 -0.1103 0.01449 0.00720 -0.0347 0.8933 0.6766
-2.000 -0.0806 0.01444 0.00732 -0.0342 0.8884 0.7493
-1.750 -0.0525 0.01456 0.00757 -0.0330 0.8822 0.8161
-1.500 -0.0151 0.01473 0.00774 -0.0337 0.8783 0.8618
-1.250 0.0265 0.01480 0.00773 -0.0358 0.8751 0.8833
-1.000 0.0607 0.01490 0.00776 -0.0368 0.8689 0.8995
-0.750 0.0974 0.01495 0.00773 -0.0384 0.8633 0.9127
-0.500 0.1364 0.01496 0.00766 -0.0404 0.8591 0.9256
-0.250 0.1788 0.01502 0.00768 -0.0432 0.8540 0.9370
0.000 0.2217 0.01506 0.00769 -0.0462 0.8482 0.9474
0.250 0.2643 0.01505 0.00765 -0.0491 0.8434 0.9579
0.500 0.3084 0.01501 0.00760 -0.0523 0.8383 0.9662
0.750 0.3478 0.01502 0.00763 -0.0549 0.8311 0.9754
1.000 0.3879 0.01496 0.00758 -0.0574 0.8255 0.9834
1.250 0.4279 0.01493 0.00759 -0.0601 0.8179 0.9910
1.500 0.4686 0.01476 0.00745 -0.0626 0.8088 0.9983
1.750 0.4919 0.01463 0.00733 -0.0615 0.7938 1.0000
2.000 0.5113 0.01448 0.00717 -0.0594 0.7769 1.0000
2.250 0.5310 0.01431 0.00698 -0.0573 0.7583 1.0000
2.500 0.5475 0.01420 0.00684 -0.0546 0.7338 1.0000
2.750 0.5655 0.01404 0.00663 -0.0520 0.7043 1.0000
3.000 0.5829 0.01398 0.00650 -0.0495 0.6724 1.0000
3.250 0.6001 0.01399 0.00645 -0.0470 0.6375 1.0000
3.500 0.6170 0.01408 0.00641 -0.0444 0.5915 1.0000
3.750 0.6314 0.01433 0.00635 -0.0413 0.5110 1.0000
4.000 0.6363 0.01525 0.00644 -0.0368 0.3668 1.0000
4.250 0.6376 0.01668 0.00701 -0.0325 0.2251 1.0000
4.500 0.6468 0.01772 0.00761 -0.0295 0.1589 1.0000
4.750 0.6597 0.01850 0.00817 -0.0270 0.1307 1.0000
5.000 0.6740 0.01919 0.00876 -0.0247 0.1141 1.0000
5.250 0.6883 0.01991 0.00937 -0.0224 0.1025 1.0000
5.500 0.7041 0.02053 0.01000 -0.0204 0.0928 1.0000
5.750 0.7193 0.02126 0.01071 -0.0182 0.0857 1.0000
6.000 0.7354 0.02194 0.01138 -0.0163 0.0791 1.0000
6.250 0.7513 0.02273 0.01215 -0.0144 0.0742 1.0000
6.500 0.7687 0.02350 0.01295 -0.0127 0.0701 1.0000
6.750 0.7860 0.02431 0.01374 -0.0111 0.0666 1.0000
7.000 0.8044 0.02523 0.01466 -0.0097 0.0633 1.0000
7.250 0.8246 0.02610 0.01559 -0.0085 0.0602 1.0000
7.500 0.8446 0.02703 0.01654 -0.0074 0.0577 1.0000
7.750 0.8652 0.02809 0.01758 -0.0065 0.0558 1.0000
8.000 0.8880 0.02939 0.01891 -0.0059 0.0542 1.0000
8.250 0.9105 0.03063 0.02033 -0.0052 0.0523 1.0000
8.500 0.9314 0.03184 0.02170 -0.0043 0.0504 1.0000
8.750 0.9510 0.03301 0.02296 -0.0034 0.0487 1.0000
9.000 0.9708 0.03435 0.02437 -0.0025 0.0475 1.0000
9.250 0.9910 0.03594 0.02602 -0.0019 0.0466 1.0000
9.500 1.0089 0.03789 0.02821 -0.0008 0.0459 1.0000
9.750 1.0228 0.04000 0.03067 0.0008 0.0451 1.0000
10.000 1.0331 0.04220 0.03322 0.0028 0.0443 1.0000
10.250 1.0404 0.04444 0.03578 0.0049 0.0434 1.0000
10.500 1.0454 0.04654 0.03817 0.0072 0.0424 1.0000
10.750 1.0483 0.04864 0.04050 0.0096 0.0417 1.0000
11.000 1.0471 0.05077 0.04284 0.0124 0.0412 1.0000
11.250 1.0417 0.05323 0.04553 0.0153 0.0409 1.0000
11.500 1.0340 0.05583 0.04835 0.0179 0.0407 1.0000
11.750 1.0229 0.05876 0.05150 0.0202 0.0405 1.0000
12.000 1.0084 0.06207 0.05503 0.0219 0.0403 1.0000
12.250 0.9877 0.06616 0.05937 0.0229 0.0403 1.0000
12.500 0.9590 0.07146 0.06494 0.0227 0.0404 1.0000
12.750 0.9171 0.07924 0.07302 0.0200 0.0407 1.0000
13.000 0.8437 0.09497 0.08906 0.0102 0.0416 1.0000
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Polar data table (+)
Polar graphs
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