NPL 9626 AIRFOIL (npl9626-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NPL 9626 AIRFOIL (npl9626-il) Reynolds number: 200,000 Max Cl/Cd: 48.48 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-npl9626-il-200000-n5.txt Download as CSV file: xf-npl9626-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NPL 9626 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.500 -0.7161 0.11800 0.11401 0.0010 1.0000 0.0178
-13.250 -0.8321 0.08342 0.07921 -0.0191 1.0000 0.0170
-13.000 -0.8749 0.07032 0.06584 -0.0291 1.0000 0.0168
-12.750 -0.9048 0.06158 0.05682 -0.0354 1.0000 0.0167
-12.500 -0.9284 0.05513 0.05010 -0.0389 1.0000 0.0167
-12.250 -0.9472 0.05021 0.04490 -0.0402 1.0000 0.0167
-12.000 -0.9618 0.04630 0.04071 -0.0399 1.0000 0.0168
-11.750 -0.9726 0.04307 0.03720 -0.0383 1.0000 0.0169
-11.500 -0.9798 0.04031 0.03415 -0.0359 1.0000 0.0171
-11.250 -0.9830 0.03783 0.03139 -0.0332 1.0000 0.0173
-11.000 -0.9802 0.03539 0.02862 -0.0309 1.0000 0.0176
-10.750 -0.9722 0.03336 0.02632 -0.0290 1.0000 0.0179
-10.500 -0.9592 0.03211 0.02499 -0.0276 1.0000 0.0182
-10.250 -0.9449 0.03090 0.02367 -0.0262 1.0000 0.0186
-10.000 -0.9299 0.02966 0.02230 -0.0248 1.0000 0.0191
-9.750 -0.9141 0.02837 0.02085 -0.0234 1.0000 0.0197
-9.500 -0.8977 0.02703 0.01931 -0.0219 1.0000 0.0205
-9.250 -0.8806 0.02582 0.01794 -0.0205 1.0000 0.0213
-9.000 -0.8628 0.02493 0.01701 -0.0192 1.0000 0.0219
-8.750 -0.8445 0.02406 0.01608 -0.0179 1.0000 0.0227
-8.500 -0.8259 0.02318 0.01509 -0.0165 1.0000 0.0238
-8.250 -0.8074 0.02233 0.01411 -0.0151 1.0000 0.0250
-8.000 -0.7893 0.02158 0.01336 -0.0137 1.0000 0.0260
-7.750 -0.7708 0.02086 0.01260 -0.0122 1.0000 0.0271
-7.500 -0.7521 0.02020 0.01183 -0.0106 1.0000 0.0287
-7.250 -0.7340 0.01952 0.01115 -0.0091 1.0000 0.0301
-7.000 -0.7152 0.01893 0.01054 -0.0076 1.0000 0.0317
-6.750 -0.6962 0.01837 0.00990 -0.0061 1.0000 0.0335
-6.500 -0.6775 0.01782 0.00938 -0.0046 1.0000 0.0354
-6.250 -0.6579 0.01737 0.00889 -0.0032 1.0000 0.0379
-6.000 -0.6387 0.01691 0.00842 -0.0017 1.0000 0.0403
-5.750 -0.6190 0.01652 0.00802 -0.0003 1.0000 0.0431
-5.500 -0.5993 0.01611 0.00760 0.0011 1.0000 0.0462
-5.250 -0.5668 0.01568 0.00715 -0.0002 0.9964 0.0508
-5.000 -0.5286 0.01517 0.00666 -0.0027 0.9902 0.0570
-4.750 -0.4891 0.01466 0.00617 -0.0054 0.9838 0.0667
-4.500 -0.4536 0.01415 0.00570 -0.0072 0.9758 0.0813
-4.250 -0.4188 0.01355 0.00524 -0.0090 0.9696 0.1093
-4.000 -0.3900 0.01274 0.00481 -0.0096 0.9606 0.1814
-3.750 -0.3599 0.01208 0.00451 -0.0105 0.9521 0.2617
-3.500 -0.3289 0.01161 0.00428 -0.0114 0.9429 0.3234
-3.250 -0.2999 0.01123 0.00409 -0.0117 0.9321 0.3761
-3.000 -0.2705 0.01081 0.00391 -0.0120 0.9217 0.4332
-2.750 -0.2422 0.01042 0.00376 -0.0120 0.9094 0.4913
-2.500 -0.2152 0.01008 0.00365 -0.0117 0.8950 0.5478
-2.250 -0.1881 0.00981 0.00354 -0.0112 0.8794 0.5973
-2.000 -0.1610 0.00960 0.00345 -0.0106 0.8626 0.6374
-1.750 -0.1339 0.00944 0.00336 -0.0100 0.8450 0.6715
-1.500 -0.1069 0.00934 0.00327 -0.0094 0.8265 0.7004
-1.250 -0.0803 0.00926 0.00321 -0.0087 0.8070 0.7274
-1.000 -0.0539 0.00921 0.00316 -0.0079 0.7859 0.7543
-0.750 -0.0275 0.00920 0.00314 -0.0071 0.7633 0.7797
-0.500 -0.0011 0.00922 0.00312 -0.0063 0.7405 0.8044
-0.250 0.0258 0.00928 0.00313 -0.0056 0.7186 0.8260
0.000 0.0532 0.00936 0.00315 -0.0051 0.6961 0.8458
0.250 0.0810 0.00946 0.00318 -0.0046 0.6734 0.8639
0.750 0.1385 0.00973 0.00327 -0.0043 0.6290 0.8954
1.000 0.1678 0.00988 0.00333 -0.0043 0.6077 0.9096
1.250 0.1984 0.01004 0.00341 -0.0046 0.5879 0.9223
1.500 0.2311 0.01021 0.00350 -0.0055 0.5695 0.9326
1.750 0.2624 0.01036 0.00359 -0.0060 0.5526 0.9432
2.000 0.2956 0.01053 0.00369 -0.0071 0.5354 0.9517
2.250 0.3286 0.01071 0.00378 -0.0081 0.5167 0.9597
2.500 0.3629 0.01091 0.00389 -0.0094 0.4977 0.9681
2.750 0.4009 0.01110 0.00402 -0.0116 0.4779 0.9749
3.000 0.4370 0.01129 0.00414 -0.0133 0.4572 0.9826
3.250 0.4749 0.01147 0.00425 -0.0156 0.4329 0.9877
3.500 0.5098 0.01170 0.00435 -0.0173 0.4012 0.9934
3.750 0.5457 0.01191 0.00446 -0.0192 0.3714 0.9977
4.000 0.5752 0.01212 0.00460 -0.0198 0.3464 1.0000
4.250 0.5973 0.01232 0.00473 -0.0188 0.3143 1.0000
4.500 0.6175 0.01274 0.00490 -0.0176 0.2589 1.0000
4.750 0.6346 0.01358 0.00528 -0.0161 0.1725 1.0000
5.000 0.6515 0.01446 0.00580 -0.0146 0.1092 1.0000
5.250 0.6706 0.01505 0.00625 -0.0132 0.0873 1.0000
5.500 0.6904 0.01553 0.00669 -0.0119 0.0761 1.0000
5.750 0.7107 0.01597 0.00711 -0.0106 0.0689 1.0000
6.000 0.7308 0.01642 0.00756 -0.0093 0.0641 1.0000
6.250 0.7509 0.01688 0.00802 -0.0080 0.0602 1.0000
6.500 0.7697 0.01744 0.00857 -0.0065 0.0571 1.0000
6.750 0.7897 0.01790 0.00909 -0.0052 0.0545 1.0000
7.000 0.8092 0.01843 0.00964 -0.0038 0.0521 1.0000
7.250 0.8276 0.01907 0.01029 -0.0024 0.0502 1.0000
7.500 0.8454 0.01979 0.01103 -0.0009 0.0485 1.0000
7.750 0.8645 0.02042 0.01173 0.0005 0.0469 1.0000
8.000 0.8833 0.02110 0.01247 0.0018 0.0452 1.0000
8.250 0.9017 0.02182 0.01320 0.0031 0.0435 1.0000
8.500 0.9179 0.02277 0.01412 0.0046 0.0418 1.0000
8.750 0.9373 0.02342 0.01489 0.0058 0.0401 1.0000
9.000 0.9559 0.02416 0.01571 0.0070 0.0385 1.0000
9.250 0.9737 0.02492 0.01651 0.0082 0.0370 1.0000
9.500 0.9904 0.02579 0.01736 0.0095 0.0358 1.0000
9.750 1.0073 0.02670 0.01837 0.0108 0.0345 1.0000
10.000 1.0240 0.02759 0.01940 0.0121 0.0333 1.0000
10.250 1.0398 0.02854 0.02044 0.0135 0.0322 1.0000
10.500 1.0547 0.02949 0.02147 0.0149 0.0314 1.0000
10.750 1.0675 0.03045 0.02245 0.0165 0.0307 1.0000
11.000 1.0792 0.03156 0.02357 0.0182 0.0300 1.0000
11.250 1.0903 0.03282 0.02505 0.0199 0.0292 1.0000
11.500 1.1004 0.03415 0.02656 0.0216 0.0283 1.0000
11.750 1.1095 0.03548 0.02803 0.0231 0.0275 1.0000
12.000 1.1178 0.03682 0.02947 0.0244 0.0269 1.0000
12.250 1.1254 0.03821 0.03096 0.0256 0.0264 1.0000
12.500 1.1324 0.03970 0.03251 0.0265 0.0259 1.0000
12.750 1.1388 0.04138 0.03424 0.0273 0.0255 1.0000
13.000 1.1400 0.04386 0.03698 0.0281 0.0250 1.0000
13.250 1.1395 0.04659 0.03995 0.0285 0.0246 1.0000
13.500 1.1373 0.04958 0.04317 0.0286 0.0242 1.0000
13.750 1.1334 0.05287 0.04666 0.0283 0.0239 1.0000
14.000 1.1277 0.05651 0.05051 0.0274 0.0236 1.0000
14.250 1.1202 0.06055 0.05474 0.0261 0.0233 1.0000
14.500 1.1105 0.06508 0.05947 0.0243 0.0231 1.0000
14.750 1.0981 0.07019 0.06478 0.0220 0.0230 1.0000
15.000 1.0824 0.07608 0.07087 0.0189 0.0229 1.0000
15.250 1.0620 0.08301 0.07801 0.0151 0.0228 1.0000
15.500 1.0342 0.09179 0.08702 0.0099 0.0228 1.0000
15.750 0.9852 0.10571 0.10127 0.0013 0.0229 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NPL 9626 AIRFOIL (npl9626-il)