RAF 89 AIRFOIL (raf89-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: RAF 89 AIRFOIL (raf89-il) Reynolds number: 200,000 Max Cl/Cd: 45.68 at α=9.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-raf89-il-200000.txt Download as CSV file: xf-raf89-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-19.750 -0.9535 0.08710 0.08087 -0.0813 1.0000 0.1071
-19.500 -0.9825 0.08067 0.07419 -0.0839 1.0000 0.1085
-19.250 -0.9919 0.07683 0.07022 -0.0850 1.0000 0.1098
-19.000 -0.9670 0.07719 0.07072 -0.0842 1.0000 0.1116
-18.750 -0.9594 0.07546 0.06901 -0.0844 1.0000 0.1132
-18.500 -0.9611 0.07266 0.06615 -0.0850 1.0000 0.1151
-18.250 -0.9714 0.06892 0.06227 -0.0859 1.0000 0.1171
-18.000 -0.9850 0.06496 0.05811 -0.0868 1.0000 0.1187
-17.750 -0.9633 0.06493 0.05826 -0.0860 1.0000 0.1211
-17.500 -0.9579 0.06316 0.05652 -0.0860 1.0000 0.1235
-17.250 -0.9593 0.06070 0.05400 -0.0861 1.0000 0.1259
-17.000 -0.9693 0.05747 0.05060 -0.0863 1.0000 0.1284
-16.750 -0.9564 0.05661 0.04987 -0.0856 1.0000 0.1313
-16.500 -0.9526 0.05491 0.04820 -0.0852 1.0000 0.1340
-16.250 -0.9558 0.05261 0.04584 -0.0848 1.0000 0.1373
-16.000 -0.9579 0.05051 0.04368 -0.0842 1.0000 0.1404
-15.750 -0.9518 0.04927 0.04256 -0.0833 1.0000 0.1438
-15.500 -0.9549 0.04735 0.04063 -0.0824 1.0000 0.1475
-15.250 -0.9627 0.04520 0.03837 -0.0812 1.0000 0.1512
-15.000 -0.9588 0.04418 0.03750 -0.0796 1.0000 0.1550
-14.750 -0.9659 0.04272 0.03606 -0.0775 1.0000 0.1587
-14.500 -0.9889 0.04173 0.03506 -0.0729 1.0000 0.1614
-14.250 -0.9730 0.04017 0.03357 -0.0748 0.9960 0.1667
-14.000 -0.9505 0.03791 0.03125 -0.0788 0.9882 0.1735
-13.750 -0.9257 0.03620 0.02956 -0.0819 0.9808 0.1799
-13.500 -0.9034 0.03427 0.02756 -0.0854 0.9745 0.1870
-13.250 -0.8804 0.03282 0.02613 -0.0877 0.9651 0.1934
-13.000 -0.8560 0.03120 0.02445 -0.0908 0.9594 0.2005
-12.750 -0.8332 0.02993 0.02317 -0.0927 0.9500 0.2070
-12.500 -0.8053 0.02868 0.02189 -0.0958 0.9438 0.2140
-12.250 -0.7710 0.02739 0.02053 -0.0998 0.9399 0.2215
-12.000 -0.7450 0.02652 0.01966 -0.1017 0.9293 0.2278
-11.750 -0.7150 0.02532 0.01829 -0.1051 0.9222 0.2352
-11.500 -0.6824 0.02472 0.01776 -0.1074 0.9132 0.2415
-11.250 -0.6599 0.02396 0.01687 -0.1086 0.9032 0.2482
-11.000 -0.6329 0.02337 0.01625 -0.1099 0.8950 0.2543
-10.750 -0.6176 0.02302 0.01588 -0.1087 0.8846 0.2595
-10.500 -0.6022 0.02254 0.01524 -0.1080 0.8764 0.2656
-10.250 -0.5930 0.02223 0.01492 -0.1054 0.8668 0.2706
-10.000 -0.5771 0.02198 0.01469 -0.1040 0.8591 0.2759
-9.750 -0.5654 0.02173 0.01431 -0.1020 0.8525 0.2817
-9.500 -0.5704 0.02154 0.01403 -0.0968 0.8443 0.2861
-9.250 -0.5476 0.02129 0.01389 -0.0962 0.8383 0.2919
-9.000 -0.5352 0.02110 0.01364 -0.0938 0.8324 0.2976
-8.750 -0.5341 0.02099 0.01342 -0.0894 0.8261 0.3026
-8.500 -0.5187 0.02077 0.01325 -0.0875 0.8203 0.3078
-8.250 -0.5052 0.02067 0.01317 -0.0851 0.8148 0.3130
-8.000 -0.4953 0.02058 0.01300 -0.0820 0.8096 0.3187
-7.750 -0.4889 0.02058 0.01286 -0.0783 0.8047 0.3238
-7.500 -0.4698 0.02039 0.01281 -0.0769 0.7995 0.3293
-7.250 -0.4579 0.02034 0.01277 -0.0741 0.7939 0.3348
-7.000 -0.4468 0.02026 0.01258 -0.0712 0.7888 0.3408
-6.750 -0.4208 0.01996 0.01229 -0.0709 0.7848 0.3471
-6.500 -0.4015 0.01990 0.01224 -0.0695 0.7804 0.3533
-6.250 -0.3945 0.01991 0.01224 -0.0658 0.7741 0.3592
-6.000 -0.3779 0.01973 0.01203 -0.0638 0.7685 0.3654
-5.750 -0.3533 0.01947 0.01183 -0.0632 0.7639 0.3721
-5.500 -0.3311 0.01932 0.01159 -0.0621 0.7599 0.3795
-5.250 -0.3143 0.01923 0.01151 -0.0602 0.7550 0.3863
-5.000 -0.2988 0.01913 0.01151 -0.0580 0.7491 0.3930
-4.750 -0.2821 0.01903 0.01137 -0.0559 0.7440 0.4007
-4.500 -0.2564 0.01874 0.01110 -0.0555 0.7398 0.4088
-4.250 -0.2293 0.01854 0.01090 -0.0553 0.7361 0.4177
-4.000 -0.2161 0.01854 0.01089 -0.0526 0.7308 0.4260
-3.750 -0.1994 0.01843 0.01093 -0.0506 0.7248 0.4344
-3.500 -0.1807 0.01830 0.01076 -0.0488 0.7198 0.4444
-3.250 -0.1532 0.01805 0.01061 -0.0486 0.7157 0.4548
-3.000 -0.1253 0.01791 0.01043 -0.0484 0.7120 0.4671
-2.750 -0.1161 0.01800 0.01068 -0.0449 0.7056 0.4768
-2.500 -0.1002 0.01798 0.01071 -0.0425 0.6996 0.4887
-2.250 -0.0768 0.01789 0.01064 -0.0414 0.6948 0.5019
-2.000 -0.0468 0.01780 0.01056 -0.0415 0.6909 0.5150
-1.750 -0.0311 0.01791 0.01067 -0.0390 0.6854 0.5283
-1.500 -0.0177 0.01803 0.01088 -0.0361 0.6785 0.5396
-1.250 0.0056 0.01797 0.01081 -0.0349 0.6733 0.5523
-1.000 0.0357 0.01790 0.01068 -0.0349 0.6691 0.5656
-0.750 0.0563 0.01797 0.01079 -0.0334 0.6635 0.5761
-0.500 0.0684 0.01805 0.01090 -0.0302 0.6561 0.5873
-0.250 0.0954 0.01791 0.01076 -0.0297 0.6508 0.5978
0.000 0.1299 0.01778 0.01057 -0.0305 0.6464 0.6092
0.250 0.1399 0.01783 0.01067 -0.0270 0.6390 0.6185
0.500 0.1624 0.01776 0.01066 -0.0257 0.6322 0.6274
0.750 0.1909 0.01759 0.01037 -0.0254 0.6270 0.6379
1.000 0.2153 0.01752 0.01038 -0.0245 0.6205 0.6458
1.250 0.2309 0.01750 0.01040 -0.0220 0.6126 0.6542
1.500 0.2581 0.01731 0.01012 -0.0216 0.6068 0.6632
1.750 0.2829 0.01725 0.01012 -0.0207 0.5999 0.6703
2.000 0.2982 0.01722 0.01012 -0.0181 0.5917 0.6781
2.250 0.3253 0.01707 0.00985 -0.0177 0.5855 0.6862
2.500 0.3429 0.01705 0.00992 -0.0157 0.5774 0.6923
2.750 0.3628 0.01700 0.00988 -0.0139 0.5695 0.6993
3.000 0.3957 0.01691 0.00965 -0.0146 0.5632 0.7068
3.250 0.3987 0.01693 0.00977 -0.0099 0.5541 0.7128
3.500 0.4247 0.01687 0.00968 -0.0093 0.5467 0.7190
3.750 0.4412 0.01692 0.00972 -0.0071 0.5390 0.7259
4.000 0.4504 0.01692 0.00968 -0.0036 0.5309 0.7328
4.250 0.4879 0.01689 0.00958 -0.0052 0.5239 0.7386
4.500 0.4840 0.01695 0.00974 0.0007 0.5157 0.7450
4.750 0.5018 0.01696 0.00970 0.0027 0.5084 0.7523
5.000 0.5211 0.01704 0.00973 0.0042 0.5014 0.7588
5.250 0.5285 0.01715 0.00993 0.0080 0.4936 0.7656
5.500 0.5567 0.01721 0.00989 0.0080 0.4867 0.7732
5.750 0.5618 0.01742 0.01013 0.0120 0.4797 0.7818
6.000 0.5759 0.01757 0.01035 0.0144 0.4725 0.7906
6.250 0.6097 0.01766 0.01031 0.0134 0.4661 0.8020
6.500 0.6095 0.01793 0.01075 0.0182 0.4591 0.8133
6.750 0.6220 0.01809 0.01094 0.0209 0.4526 0.8264
7.000 0.6628 0.01822 0.01100 0.0184 0.4460 0.8364
7.250 0.6644 0.01853 0.01147 0.0226 0.4396 0.8531
7.500 0.6924 0.01878 0.01178 0.0221 0.4326 0.8699
7.750 0.7416 0.01910 0.01208 0.0177 0.4257 0.8845
8.000 0.7770 0.01971 0.01284 0.0150 0.4180 0.9040
8.250 0.8298 0.02014 0.01322 0.0098 0.4106 0.9196
8.500 0.8791 0.02086 0.01399 0.0047 0.4032 0.9320
8.750 0.9231 0.02150 0.01466 0.0005 0.3956 0.9447
9.000 0.9755 0.02193 0.01494 -0.0041 0.3891 0.9564
9.250 1.0038 0.02281 0.01596 -0.0060 0.3822 0.9717
9.500 1.0445 0.02342 0.01656 -0.0095 0.3753 0.9828
9.750 1.0936 0.02394 0.01695 -0.0139 0.3690 0.9908
10.000 1.1161 0.02473 0.01787 -0.0152 0.3626 1.0000
10.250 1.1051 0.02498 0.01807 -0.0097 0.3585 1.0000
10.500 1.1165 0.02524 0.01822 -0.0074 0.3539 1.0000
10.750 1.1238 0.02575 0.01870 -0.0047 0.3492 1.0000
11.000 1.1167 0.02651 0.01953 -0.0003 0.3445 1.0000
11.250 1.1222 0.02707 0.02008 0.0024 0.3396 1.0000
11.500 1.1432 0.02736 0.02025 0.0032 0.3348 1.0000
11.750 1.1451 0.02821 0.02115 0.0061 0.3302 1.0000
12.000 1.1420 0.02918 0.02219 0.0094 0.3251 1.0000
12.250 1.1507 0.02980 0.02277 0.0113 0.3199 1.0000
12.500 1.1698 0.03020 0.02306 0.0122 0.3144 1.0000
12.750 1.1589 0.03164 0.02463 0.0157 0.3092 1.0000
13.000 1.1626 0.03257 0.02557 0.0178 0.3036 1.0000
13.250 1.1876 0.03274 0.02558 0.0180 0.2984 1.0000
13.500 1.1782 0.03444 0.02743 0.0208 0.2938 1.0000
13.750 1.1792 0.03574 0.02879 0.0227 0.2887 1.0000
14.000 1.1914 0.03649 0.02949 0.0237 0.2840 1.0000
14.250 1.2047 0.03731 0.03027 0.0246 0.2794 1.0000
14.500 1.1983 0.03922 0.03233 0.0266 0.2747 1.0000
14.750 1.2035 0.04049 0.03361 0.0278 0.2698 1.0000
15.000 1.2269 0.04069 0.03366 0.0280 0.2652 1.0000
15.250 1.2168 0.04307 0.03621 0.0298 0.2606 1.0000
15.500 1.2155 0.04496 0.03817 0.0309 0.2556 1.0000
15.750 1.2273 0.04591 0.03908 0.0316 0.2511 1.0000
16.000 1.2337 0.04737 0.04054 0.0323 0.2464 1.0000
16.250 1.2249 0.05004 0.04334 0.0334 0.2414 1.0000
16.500 1.2311 0.05153 0.04482 0.0339 0.2365 1.0000
16.750 1.2435 0.05257 0.04581 0.0344 0.2319 1.0000
17.000 1.2294 0.05599 0.04940 0.0351 0.2266 1.0000
17.250 1.2329 0.05786 0.05126 0.0354 0.2216 1.0000
17.500 1.2430 0.05915 0.05250 0.0357 0.2168 1.0000
17.750 1.2278 0.06300 0.05651 0.0359 0.2114 1.0000
18.000 1.2332 0.06482 0.05829 0.0360 0.2064 1.0000
18.250 1.2307 0.06751 0.06101 0.0361 0.2010 1.0000
18.500 1.2203 0.07115 0.06476 0.0359 0.1955 1.0000
18.750 1.2376 0.07169 0.06514 0.0360 0.1904 1.0000
19.000 1.2128 0.07715 0.07082 0.0354 0.1846 1.0000
19.250 1.2139 0.07966 0.07330 0.0350 0.1794 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RAF 89 AIRFOIL (raf89-il)