RAF 89 AIRFOIL (raf89-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAF 89 AIRFOIL (raf89-il) Reynolds number: 50,000 Max Cl/Cd: 6.99 at α=8° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-raf89-il-50000-n5.txt Download as CSV file: xf-raf89-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAF 89 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
------ -------- --------- --------- -------- -------- --------
-13.750 -0.3891 0.11960 0.11058 -0.0560 1.0000 0.1877
-13.500 -0.3849 0.12054 0.11162 -0.0529 1.0000 0.1899
-13.250 -0.3934 0.12033 0.11150 -0.0499 1.0000 0.1930
-13.000 -0.4287 0.11235 0.10349 -0.0528 0.9970 0.2022
-12.500 -0.4191 0.10330 0.09433 -0.0601 0.9819 0.2167
-12.250 -0.3751 0.10431 0.09531 -0.0619 0.9755 0.2196
-11.250 -0.3654 0.08817 0.07893 -0.0730 0.9475 0.2504
-10.750 -0.3679 0.07988 0.07050 -0.0779 0.9344 0.2671
-10.250 -0.3684 0.07138 0.06182 -0.0835 0.9236 0.2848
-9.750 -0.4076 0.06195 0.05220 -0.0855 0.9075 0.3033
-9.500 -0.3380 0.06467 0.05493 -0.0875 0.9046 0.3058
-9.250 -0.3056 0.06487 0.05511 -0.0882 0.8989 0.3103
-9.000 -0.3985 0.05540 0.04543 -0.0865 0.8870 0.3245
-8.750 -0.3379 0.05690 0.04695 -0.0888 0.8839 0.3287
-8.500 -0.4673 0.04928 0.03902 -0.0795 0.8683 0.3395
-8.250 -0.4293 0.04929 0.03905 -0.0809 0.8639 0.3452
-8.000 -0.3803 0.04961 0.03938 -0.0834 0.8606 0.3514
-7.750 -0.4511 0.04716 0.03677 -0.0735 0.8479 0.3576
-7.500 -0.4888 0.04550 0.03495 -0.0662 0.8391 0.3634
-7.250 -0.4190 0.04593 0.03547 -0.0717 0.8371 0.3700
-7.000 -0.4479 0.04545 0.03494 -0.0643 0.8276 0.3751
-6.750 -0.4922 0.04441 0.03374 -0.0548 0.8174 0.3806
-6.500 -0.4574 0.04405 0.03340 -0.0561 0.8136 0.3876
-6.250 -0.4178 0.04373 0.03308 -0.0581 0.8107 0.3954
-6.000 -0.4773 0.04376 0.03306 -0.0453 0.7970 0.3986
-5.750 -0.4716 0.04296 0.03215 -0.0426 0.7913 0.4063
-5.500 -0.4202 0.04299 0.03226 -0.0461 0.7884 0.4141
-5.250 -0.4586 0.04312 0.03237 -0.0362 0.7768 0.4184
-5.000 -0.4630 0.04247 0.03160 -0.0318 0.7697 0.4263
-4.750 -0.4135 0.04260 0.03183 -0.0348 0.7662 0.4342
-4.500 -0.4135 0.04235 0.03152 -0.0307 0.7593 0.4421
-4.250 -0.4336 0.04245 0.03162 -0.0237 0.7486 0.4479
-4.000 -0.3960 0.04252 0.03175 -0.0249 0.7441 0.4560
-3.750 -0.3701 0.04190 0.03104 -0.0247 0.7403 0.4670
-3.500 -0.3952 0.04273 0.03195 -0.0168 0.7279 0.4711
-3.250 -0.3751 0.04265 0.03186 -0.0155 0.7218 0.4805
-3.000 -0.3408 0.04241 0.03162 -0.0161 0.7180 0.4906
-2.750 -0.3581 0.04299 0.03223 -0.0097 0.7070 0.4975
-2.500 -0.3488 0.04293 0.03212 -0.0069 0.6997 0.5077
-2.250 -0.3161 0.04276 0.03194 -0.0071 0.6955 0.5187
-2.000 -0.3267 0.04315 0.03232 -0.0018 0.6854 0.5276
-1.750 -0.3161 0.04337 0.03254 0.0008 0.6775 0.5379
-1.500 -0.2877 0.04313 0.03227 0.0012 0.6731 0.5500
-1.250 -0.2947 0.04354 0.03264 0.0060 0.6631 0.5607
-1.000 -0.2822 0.04375 0.03287 0.0083 0.6552 0.5711
-0.750 -0.2599 0.04333 0.03234 0.0095 0.6509 0.5860
-0.500 -0.2587 0.04414 0.03324 0.0130 0.6403 0.5942
-0.250 -0.2495 0.04406 0.03306 0.0157 0.6330 0.6080
0.000 -0.2112 0.04400 0.03308 0.0149 0.6291 0.6189
0.250 -0.2267 0.04471 0.03373 0.0203 0.6173 0.6308
0.500 -0.1978 0.04489 0.03398 0.0205 0.6114 0.6411
0.750 -0.1691 0.04450 0.03353 0.0210 0.6077 0.6549
1.000 -0.1768 0.04576 0.03486 0.0250 0.5952 0.6634
1.250 -0.1544 0.04558 0.03464 0.0262 0.5900 0.6769
1.500 -0.1155 0.04540 0.03449 0.0254 0.5867 0.6877
1.750 -0.1351 0.04668 0.03575 0.0305 0.5735 0.6988
2.000 -0.1012 0.04667 0.03579 0.0302 0.5691 0.7093
2.250 -0.0698 0.04621 0.03527 0.0305 0.5662 0.7223
2.500 -0.0827 0.04795 0.03708 0.0342 0.5524 0.7310
2.750 -0.0561 0.04766 0.03675 0.0349 0.5486 0.7441
3.250 -0.0407 0.04920 0.03832 0.0389 0.5319 0.7660
3.500 -0.0037 0.04900 0.03815 0.0382 0.5284 0.7768
3.750 -0.0178 0.05070 0.03987 0.0417 0.5166 0.7886
4.000 0.0143 0.05085 0.04007 0.0411 0.5117 0.7996
4.250 0.0465 0.05051 0.03972 0.0410 0.5086 0.8127
4.500 0.0412 0.05282 0.04213 0.0425 0.4962 0.8236
4.750 0.0727 0.05290 0.04223 0.0418 0.4918 0.8369
5.000 0.1174 0.05270 0.04207 0.0398 0.4889 0.8502
5.250 0.1152 0.05559 0.04507 0.0398 0.4760 0.8632
5.500 0.1552 0.05580 0.04532 0.0378 0.4719 0.8790
5.750 0.2138 0.05567 0.04523 0.0335 0.4692 0.8896
6.250 0.2602 0.05949 0.04917 0.0286 0.4519 0.9117
6.500 0.3095 0.05911 0.04878 0.0256 0.4495 0.9227
7.000 0.3509 0.06366 0.05343 0.0202 0.4320 0.9446
7.250 0.3998 0.06321 0.05297 0.0172 0.4297 0.9567
7.750 0.4338 0.06864 0.05850 0.0114 0.4120 0.9871
8.000 0.4767 0.06822 0.05803 0.0090 0.4100 1.0000
8.500 0.4258 0.07353 0.06317 0.0155 0.3926 1.0000
8.750 0.3965 0.07783 0.06743 0.0174 0.3823 1.0000
9.000 0.4005 0.07929 0.06883 0.0187 0.3767 1.0000
9.250 0.4198 0.07942 0.06889 0.0197 0.3735 1.0000
9.500 0.4442 0.07911 0.06850 0.0207 0.3713 1.0000
9.750 0.4022 0.08565 0.07506 0.0216 0.3584 1.0000
10.000 0.4204 0.08616 0.07552 0.0224 0.3549 1.0000
10.250 0.4454 0.08601 0.07532 0.0232 0.3526 1.0000
10.500 0.4080 0.09250 0.08185 0.0233 0.3410 1.0000
10.750 0.4231 0.09349 0.08280 0.0239 0.3370 1.0000
11.000 0.4459 0.09370 0.08297 0.0246 0.3344 1.0000
11.250 0.4168 0.09960 0.08891 0.0244 0.3244 1.0000
11.500 0.4282 0.10111 0.09041 0.0248 0.3199 1.0000
11.750 0.4487 0.10163 0.09090 0.0253 0.3169 1.0000
12.250 0.4358 0.10881 0.09812 0.0250 0.3034 1.0000
12.500 0.4539 0.10969 0.09899 0.0253 0.3001 1.0000
12.750 0.4692 0.11091 0.10018 0.0256 0.2967 1.0000
13.000 0.4465 0.11647 0.10580 0.0247 0.2876 1.0000
13.250 0.4619 0.11767 0.10700 0.0249 0.2838 1.0000
13.500 0.4843 0.11809 0.10741 0.0253 0.2812 1.0000
13.750 0.4610 0.12388 0.11324 0.0240 0.2722 1.0000
14.000 0.4734 0.12542 0.11479 0.0240 0.2678 1.0000
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