AIRFOIL PROFILE12A 9.00% (rg12a-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: AIRFOIL PROFILE12A 9.00% (rg12a-il) Reynolds number: 100,000 Max Cl/Cd: 49.63 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rg12a-il-100000-n5.txt Download as CSV file: xf-rg12a-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: AIRFOIL PROFILE12A 9.00%
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 -0.4810 0.11314 0.10783 -0.0198 1.0000 0.0284
-10.250 -0.4784 0.10923 0.10396 -0.0205 1.0000 0.0267
-10.000 -0.4781 0.10500 0.09977 -0.0219 1.0000 0.0253
-9.500 -0.4883 0.09268 0.08758 -0.0292 1.0000 0.0212
-9.250 -0.4893 0.08890 0.08385 -0.0300 1.0000 0.0208
-9.000 -0.4926 0.08457 0.07959 -0.0317 1.0000 0.0206
-8.750 -0.4967 0.08033 0.07541 -0.0335 1.0000 0.0203
-8.500 -0.5047 0.07553 0.07067 -0.0362 1.0000 0.0201
-8.250 -0.5162 0.07114 0.06634 -0.0383 1.0000 0.0198
-8.000 -0.5252 0.06658 0.06176 -0.0398 1.0000 0.0196
-7.750 -0.5329 0.06193 0.05704 -0.0405 1.0000 0.0193
-7.500 -0.5380 0.05742 0.05241 -0.0407 1.0000 0.0191
-7.250 -0.5406 0.05291 0.04773 -0.0403 1.0000 0.0188
-7.000 -0.5400 0.04852 0.04308 -0.0395 1.0000 0.0185
-6.750 -0.5360 0.04429 0.03854 -0.0384 1.0000 0.0183
-6.500 -0.5290 0.04015 0.03399 -0.0372 1.0000 0.0183
-6.250 -0.5182 0.03644 0.02984 -0.0358 1.0000 0.0184
-6.000 -0.5038 0.03331 0.02622 -0.0345 1.0000 0.0193
-5.750 -0.4866 0.03059 0.02293 -0.0331 1.0000 0.0204
-5.500 -0.4672 0.02843 0.02023 -0.0318 1.0000 0.0213
-5.250 -0.4478 0.02577 0.01727 -0.0306 1.0000 0.0219
-5.000 -0.4270 0.02390 0.01519 -0.0296 1.0000 0.0228
-4.750 -0.4056 0.02242 0.01354 -0.0285 1.0000 0.0241
-4.500 -0.3838 0.02111 0.01206 -0.0273 1.0000 0.0257
-4.250 -0.3618 0.02009 0.01088 -0.0262 1.0000 0.0288
-4.000 -0.3404 0.01908 0.00981 -0.0253 1.0000 0.0335
-3.750 -0.3183 0.01820 0.00882 -0.0243 1.0000 0.0390
-3.500 -0.2960 0.01740 0.00800 -0.0235 1.0000 0.0515
-3.250 -0.2736 0.01655 0.00725 -0.0228 1.0000 0.0757
-3.000 -0.2379 0.01565 0.00678 -0.0251 0.9947 0.1535
-2.750 -0.2042 0.01467 0.00639 -0.0271 0.9886 0.2920
-2.500 -0.1717 0.01362 0.00622 -0.0286 0.9829 0.5018
-2.250 -0.1480 0.01321 0.00657 -0.0266 0.9741 0.7211
-2.000 -0.1192 0.01325 0.00661 -0.0260 0.9650 0.7949
-1.750 -0.0880 0.01328 0.00656 -0.0258 0.9568 0.8401
-1.500 -0.0595 0.01325 0.00645 -0.0252 0.9467 0.8744
-1.250 -0.0277 0.01321 0.00634 -0.0252 0.9375 0.9043
-1.000 0.0143 0.01318 0.00618 -0.0274 0.9311 0.9312
-0.750 0.0608 0.01315 0.00604 -0.0307 0.9235 0.9546
-0.500 0.1170 0.01309 0.00587 -0.0362 0.9184 0.9724
-0.250 0.1697 0.01299 0.00567 -0.0412 0.9102 0.9841
0.000 0.2197 0.01283 0.00544 -0.0458 0.9009 0.9906
0.250 0.2643 0.01267 0.00523 -0.0492 0.8883 0.9967
0.500 0.3020 0.01253 0.00504 -0.0513 0.8723 1.0000
0.750 0.3334 0.01240 0.00487 -0.0520 0.8535 1.0000
1.000 0.3659 0.01228 0.00471 -0.0529 0.8344 1.0000
1.250 0.3952 0.01221 0.00461 -0.0531 0.8135 1.0000
1.500 0.4235 0.01217 0.00452 -0.0531 0.7913 1.0000
1.750 0.4491 0.01218 0.00448 -0.0525 0.7676 1.0000
2.000 0.4746 0.01223 0.00446 -0.0519 0.7440 1.0000
2.250 0.4992 0.01231 0.00450 -0.0512 0.7199 1.0000
2.500 0.5227 0.01244 0.00457 -0.0502 0.6950 1.0000
2.750 0.5464 0.01260 0.00466 -0.0493 0.6701 1.0000
3.000 0.5702 0.01278 0.00478 -0.0484 0.6448 1.0000
3.250 0.5939 0.01299 0.00496 -0.0475 0.6187 1.0000
3.500 0.6175 0.01323 0.00514 -0.0465 0.5918 1.0000
3.750 0.6410 0.01349 0.00535 -0.0456 0.5640 1.0000
4.000 0.6644 0.01377 0.00558 -0.0447 0.5353 1.0000
4.250 0.6875 0.01408 0.00587 -0.0438 0.5053 1.0000
4.500 0.7104 0.01442 0.00616 -0.0428 0.4735 1.0000
4.750 0.7331 0.01479 0.00648 -0.0418 0.4408 1.0000
5.000 0.7553 0.01522 0.00683 -0.0408 0.4075 1.0000
5.500 0.7990 0.01619 0.00770 -0.0388 0.3360 1.0000
5.750 0.8203 0.01675 0.00820 -0.0378 0.2980 1.0000
6.000 0.8412 0.01738 0.00875 -0.0367 0.2602 1.0000
6.250 0.8616 0.01809 0.00937 -0.0357 0.2220 1.0000
6.500 0.8809 0.01895 0.01005 -0.0346 0.1768 1.0000
6.750 0.8987 0.02002 0.01086 -0.0335 0.1276 1.0000
7.000 0.9173 0.02106 0.01176 -0.0324 0.0934 1.0000
7.250 0.9353 0.02221 0.01286 -0.0311 0.0696 1.0000
7.500 0.9514 0.02356 0.01413 -0.0296 0.0531 1.0000
7.750 0.9682 0.02481 0.01543 -0.0281 0.0429 1.0000
8.000 0.9835 0.02618 0.01692 -0.0264 0.0383 1.0000
8.250 0.9991 0.02754 0.01848 -0.0248 0.0350 1.0000
8.500 1.0136 0.02900 0.02006 -0.0231 0.0327 1.0000
8.750 1.0261 0.03089 0.02202 -0.0213 0.0309 1.0000
9.000 1.0425 0.03263 0.02395 -0.0199 0.0295 1.0000
9.250 1.0588 0.03428 0.02586 -0.0185 0.0274 1.0000
9.500 1.0733 0.03603 0.02782 -0.0171 0.0254 1.0000
9.750 1.0867 0.03812 0.03012 -0.0157 0.0244 1.0000
10.000 1.0982 0.04039 0.03259 -0.0142 0.0235 1.0000
10.250 1.1071 0.04299 0.03540 -0.0126 0.0228 1.0000
10.500 1.1111 0.04627 0.03892 -0.0107 0.0221 1.0000
10.750 1.1089 0.04913 0.04216 -0.0081 0.0216 1.0000
11.000 1.1025 0.05178 0.04518 -0.0056 0.0210 1.0000
11.250 1.0928 0.05489 0.04863 -0.0035 0.0206 1.0000
11.500 1.0802 0.05836 0.05241 -0.0023 0.0200 1.0000
11.750 1.0659 0.06242 0.05673 -0.0022 0.0199 1.0000
12.000 1.0488 0.06716 0.06174 -0.0032 0.0197 1.0000
12.250 1.0306 0.07267 0.06747 -0.0054 0.0198 1.0000
12.500 1.0100 0.07920 0.07422 -0.0091 0.0199 1.0000
12.750 0.9890 0.08668 0.08187 -0.0139 0.0202 1.0000
13.000 0.9670 0.09525 0.09054 -0.0197 0.0205 1.0000
13.250 0.9463 0.10441 0.09979 -0.0258 0.0211 1.0000
13.500 0.9257 0.11438 0.10982 -0.0322 0.0215 1.0000
13.750 0.9083 0.12413 0.11959 -0.0379 0.0219 1.0000
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Polar data table (+)
Polar graphs
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