AH 85-L-120/17 (ah85l120-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: AH 85-L-120/17 (ah85l120-il) Reynolds number: 100,000 Max Cl/Cd: 29.76 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ah85l120-il-100000-n5.txt Download as CSV file: xf-ah85l120-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: AH 85-L-120/17
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
------ -------- --------- --------- -------- -------- --------
-11.750 -0.6175 0.09111 0.08619 -0.0392 1.0000 0.0229
-11.500 -0.6399 0.08192 0.07694 -0.0454 1.0000 0.0221
-11.250 -0.6638 0.07501 0.06993 -0.0492 1.0000 0.0218
-11.000 -0.6884 0.06938 0.06416 -0.0509 1.0000 0.0216
-10.750 -0.7109 0.06479 0.05940 -0.0508 1.0000 0.0214
-10.500 -0.7310 0.06095 0.05537 -0.0494 1.0000 0.0211
-10.250 -0.7490 0.05769 0.05192 -0.0469 1.0000 0.0210
-10.000 -0.7666 0.05480 0.04883 -0.0432 1.0000 0.0210
-9.750 -0.7833 0.05216 0.04593 -0.0386 1.0000 0.0207
-9.500 -0.7951 0.04977 0.04332 -0.0341 1.0000 0.0208
-9.250 -0.8067 0.04776 0.04109 -0.0291 1.0000 0.0209
-9.000 -0.8187 0.04594 0.03902 -0.0236 1.0000 0.0206
-8.750 -0.8277 0.04420 0.03707 -0.0184 1.0000 0.0210
-8.500 -0.8305 0.04219 0.03477 -0.0141 0.9994 0.0211
-8.250 -0.8155 0.03946 0.03164 -0.0132 0.9945 0.0214
-8.000 -0.7961 0.03681 0.02859 -0.0127 0.9907 0.0220
-7.750 -0.7750 0.03501 0.02647 -0.0124 0.9868 0.0244
-7.500 -0.7503 0.03298 0.02409 -0.0124 0.9838 0.0255
-7.250 -0.7223 0.03099 0.02170 -0.0127 0.9818 0.0266
-7.000 -0.6951 0.02927 0.01972 -0.0127 0.9796 0.0275
-6.750 -0.6689 0.02726 0.01762 -0.0128 0.9775 0.0286
-6.500 -0.6450 0.02580 0.01613 -0.0128 0.9746 0.0304
-6.250 -0.6202 0.02473 0.01498 -0.0129 0.9718 0.0347
-6.000 -0.5953 0.02386 0.01393 -0.0129 0.9693 0.0384
-5.750 -0.5801 0.02276 0.01281 -0.0112 0.9650 0.0424
-5.500 -0.5581 0.02195 0.01184 -0.0107 0.9615 0.0497
-5.250 -0.5354 0.02096 0.01093 -0.0104 0.9587 0.0666
-5.000 -0.5156 0.01983 0.01013 -0.0097 0.9558 0.1193
-4.750 -0.5095 0.01845 0.00959 -0.0067 0.9505 0.2620
-4.500 -0.5031 0.01697 0.00920 -0.0034 0.9462 0.4598
-4.250 -0.4723 0.01729 0.01059 -0.0015 0.9456 0.6804
-4.000 -0.4450 0.01818 0.01142 0.0001 0.9433 0.7454
-3.750 -0.4237 0.01879 0.01188 0.0022 0.9394 0.7669
-3.500 -0.4020 0.01920 0.01212 0.0039 0.9355 0.7847
-3.250 -0.3735 0.01962 0.01236 0.0043 0.9327 0.8003
-3.000 -0.3371 0.02010 0.01265 0.0032 0.9311 0.8122
-2.750 -0.3003 0.02046 0.01286 0.0018 0.9296 0.8225
-2.500 -0.2825 0.02062 0.01292 0.0038 0.9249 0.8332
-2.250 -0.2534 0.02073 0.01289 0.0035 0.9216 0.8402
-2.000 -0.2253 0.02073 0.01278 0.0032 0.9183 0.8472
-1.750 -0.1874 0.02078 0.01273 0.0010 0.9165 0.8515
-1.500 -0.1739 0.02075 0.01263 0.0034 0.9109 0.8600
-1.250 -0.1406 0.02080 0.01261 0.0021 0.9076 0.8637
-1.000 -0.1101 0.02079 0.01253 0.0012 0.9044 0.8689
-0.750 -0.0798 0.02073 0.01243 0.0003 0.9015 0.8743
-0.500 -0.0550 0.02080 0.01247 0.0006 0.8962 0.8784
-0.250 -0.0297 0.02078 0.01243 0.0007 0.8916 0.8837
0.000 0.0001 0.02073 0.01237 0.0000 0.8883 0.8883
0.250 0.0298 0.02078 0.01243 -0.0007 0.8837 0.8916
0.500 0.0549 0.02080 0.01247 -0.0006 0.8784 0.8963
0.750 0.0801 0.02073 0.01243 -0.0004 0.8743 0.9015
1.000 0.1098 0.02079 0.01254 -0.0011 0.8689 0.9043
1.250 0.1404 0.02081 0.01261 -0.0020 0.8637 0.9076
1.500 0.1737 0.02076 0.01264 -0.0034 0.8601 0.9110
1.750 0.1877 0.02079 0.01273 -0.0010 0.8515 0.9165
2.000 0.2252 0.02073 0.01278 -0.0031 0.8472 0.9183
2.250 0.2532 0.02073 0.01288 -0.0035 0.8400 0.9215
2.500 0.2828 0.02061 0.01290 -0.0038 0.8332 0.9248
2.750 0.3006 0.02046 0.01286 -0.0019 0.8226 0.9296
3.000 0.3377 0.02011 0.01266 -0.0033 0.8125 0.9311
3.250 0.3741 0.01962 0.01233 -0.0043 0.8005 0.9327
3.500 0.4026 0.01920 0.01212 -0.0040 0.7851 0.9355
3.750 0.4237 0.01877 0.01185 -0.0022 0.7664 0.9394
4.000 0.4444 0.01819 0.01143 0.0000 0.7446 0.9434
4.250 0.4722 0.01734 0.01067 0.0014 0.6851 0.9456
4.500 0.5039 0.01693 0.00921 0.0033 0.4661 0.9462
4.750 0.5097 0.01845 0.00959 0.0066 0.2615 0.9506
5.000 0.5166 0.01977 0.01012 0.0096 0.1251 0.9557
5.250 0.5348 0.02106 0.01097 0.0105 0.0636 0.9587
5.500 0.5584 0.02194 0.01183 0.0106 0.0488 0.9615
5.750 0.5802 0.02279 0.01280 0.0113 0.0418 0.9650
6.000 0.5955 0.02387 0.01396 0.0130 0.0379 0.9693
6.250 0.6203 0.02473 0.01499 0.0129 0.0330 0.9718
6.500 0.6450 0.02582 0.01614 0.0128 0.0303 0.9746
6.750 0.6691 0.02726 0.01763 0.0129 0.0287 0.9774
7.000 0.6954 0.02931 0.01978 0.0128 0.0272 0.9796
7.250 0.7225 0.03102 0.02174 0.0127 0.0266 0.9818
7.500 0.7504 0.03296 0.02407 0.0124 0.0253 0.9839
7.750 0.7748 0.03486 0.02633 0.0125 0.0236 0.9870
8.000 0.7960 0.03680 0.02856 0.0127 0.0218 0.9907
8.250 0.8153 0.03939 0.03154 0.0132 0.0211 0.9945
8.500 0.8304 0.04219 0.03474 0.0141 0.0209 0.9995
8.750 0.8275 0.04414 0.03695 0.0183 0.0206 1.0000
9.000 0.8185 0.04593 0.03899 0.0236 0.0205 1.0000
9.250 0.8068 0.04775 0.04107 0.0291 0.0207 1.0000
9.500 0.7955 0.04972 0.04326 0.0340 0.0207 1.0000
9.750 0.7825 0.05227 0.04604 0.0387 0.0207 1.0000
10.000 0.7682 0.05468 0.04864 0.0430 0.0207 1.0000
10.250 0.7502 0.05758 0.05179 0.0468 0.0209 1.0000
10.500 0.7319 0.06087 0.05527 0.0493 0.0210 1.0000
10.750 0.7096 0.06498 0.05958 0.0508 0.0213 1.0000
11.000 0.6888 0.06937 0.06413 0.0508 0.0215 1.0000
11.250 0.6631 0.07519 0.07010 0.0491 0.0217 1.0000
11.500 0.6394 0.08217 0.07719 0.0452 0.0222 1.0000
11.750 0.6173 0.09137 0.08643 0.0389 0.0226 1.0000
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Polar data table (+)
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