NPL 9626 AIRFOIL (npl9626-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NPL 9626 AIRFOIL (npl9626-il) Reynolds number: 50,000 Max Cl/Cd: 29.62 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-npl9626-il-50000-n5.txt Download as CSV file: xf-npl9626-il-50000-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 = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.6626 0.09285 0.08516 -0.0215 1.0000 0.0566
-10.750 -0.6979 0.08201 0.07424 -0.0299 1.0000 0.0550
-10.500 -0.7408 0.07346 0.06549 -0.0352 1.0000 0.0536
-10.250 -0.7725 0.06830 0.06009 -0.0352 1.0000 0.0530
-10.000 -0.7838 0.06479 0.05641 -0.0341 1.0000 0.0535
-9.750 -0.7893 0.06150 0.05294 -0.0329 1.0000 0.0542
-9.500 -0.7927 0.05823 0.04942 -0.0315 1.0000 0.0552
-9.250 -0.7939 0.05492 0.04579 -0.0298 1.0000 0.0562
-9.000 -0.7922 0.05162 0.04211 -0.0281 1.0000 0.0573
-8.750 -0.7866 0.04840 0.03846 -0.0263 1.0000 0.0582
-8.500 -0.7776 0.04537 0.03495 -0.0245 1.0000 0.0593
-8.250 -0.7647 0.04261 0.03176 -0.0228 1.0000 0.0607
-8.000 -0.7468 0.04055 0.02966 -0.0219 1.0000 0.0629
-7.750 -0.7289 0.03868 0.02760 -0.0207 1.0000 0.0660
-7.500 -0.7108 0.03681 0.02530 -0.0192 1.0000 0.0699
-7.250 -0.6909 0.03513 0.02369 -0.0183 1.0000 0.0739
-7.000 -0.6705 0.03366 0.02203 -0.0171 1.0000 0.0792
-6.750 -0.6491 0.03210 0.02041 -0.0161 1.0000 0.0838
-6.500 -0.6278 0.03074 0.01901 -0.0149 1.0000 0.0888
-6.250 -0.6069 0.02948 0.01765 -0.0136 1.0000 0.0951
-6.000 -0.5878 0.02826 0.01644 -0.0121 1.0000 0.1019
-5.750 -0.5692 0.02709 0.01524 -0.0105 1.0000 0.1100
-5.500 -0.5514 0.02597 0.01414 -0.0089 1.0000 0.1222
-5.250 -0.5346 0.02476 0.01309 -0.0073 1.0000 0.1393
-5.000 -0.5183 0.02346 0.01209 -0.0057 1.0000 0.1733
-4.750 -0.5041 0.02206 0.01134 -0.0040 1.0000 0.2592
-4.500 -0.4903 0.02110 0.01093 -0.0018 1.0000 0.3658
-4.250 -0.4753 0.02051 0.01064 0.0006 1.0000 0.4418
-4.000 -0.4590 0.02008 0.01043 0.0031 1.0000 0.5010
-3.750 -0.4423 0.01977 0.01032 0.0056 1.0000 0.5580
-3.500 -0.4268 0.01957 0.01043 0.0088 1.0000 0.6246
-3.250 -0.4100 0.01950 0.01055 0.0120 1.0000 0.6836
-3.000 -0.3896 0.01949 0.01061 0.0143 1.0000 0.7262
-2.750 -0.3671 0.01947 0.01058 0.0160 1.0000 0.7602
-2.500 -0.3431 0.01951 0.01059 0.0172 1.0000 0.7939
-2.250 -0.3163 0.01962 0.01063 0.0179 1.0000 0.8274
-2.000 -0.2859 0.01978 0.01071 0.0177 1.0000 0.8601
-1.750 -0.2518 0.01998 0.01081 0.0166 1.0000 0.8916
-1.500 -0.2015 0.02022 0.01090 0.0121 0.9959 0.9175
-1.250 -0.1386 0.02042 0.01093 0.0050 0.9852 0.9398
-1.000 -0.0719 0.02051 0.01088 -0.0030 0.9743 0.9569
-0.750 -0.0049 0.02048 0.01074 -0.0112 0.9622 0.9720
-0.500 0.0615 0.02034 0.01053 -0.0193 0.9489 0.9853
-0.250 0.1254 0.02010 0.01025 -0.0269 0.9324 0.9974
0.000 0.1706 0.01988 0.00999 -0.0309 0.9105 1.0000
0.250 0.2058 0.01969 0.00977 -0.0328 0.8852 1.0000
0.500 0.2387 0.01952 0.00956 -0.0339 0.8599 1.0000
0.750 0.2675 0.01941 0.00941 -0.0341 0.8343 1.0000
1.000 0.2941 0.01934 0.00929 -0.0338 0.8094 1.0000
1.250 0.3189 0.01933 0.00921 -0.0330 0.7855 1.0000
1.500 0.3418 0.01938 0.00921 -0.0319 0.7618 1.0000
1.750 0.3648 0.01945 0.00923 -0.0308 0.7402 1.0000
2.000 0.3876 0.01956 0.00927 -0.0296 0.7200 1.0000
2.250 0.4094 0.01973 0.00941 -0.0284 0.6998 1.0000
2.500 0.4315 0.01994 0.00960 -0.0272 0.6809 1.0000
2.750 0.4536 0.02017 0.00981 -0.0260 0.6633 1.0000
3.000 0.4760 0.02043 0.01006 -0.0249 0.6469 1.0000
3.250 0.4988 0.02071 0.01036 -0.0238 0.6314 1.0000
3.500 0.5223 0.02100 0.01067 -0.0229 0.6167 1.0000
3.750 0.5449 0.02135 0.01108 -0.0219 0.6014 1.0000
4.000 0.5675 0.02171 0.01152 -0.0209 0.5865 1.0000
4.250 0.5897 0.02206 0.01194 -0.0196 0.5704 1.0000
4.500 0.6093 0.02219 0.01202 -0.0175 0.5444 1.0000
4.750 0.6251 0.02223 0.01192 -0.0146 0.5065 1.0000
5.000 0.6399 0.02231 0.01181 -0.0116 0.4638 1.0000
5.250 0.6566 0.02257 0.01201 -0.0093 0.4279 1.0000
5.500 0.6742 0.02291 0.01235 -0.0074 0.3941 1.0000
5.750 0.6907 0.02332 0.01278 -0.0053 0.3496 1.0000
6.000 0.7049 0.02393 0.01320 -0.0030 0.2773 1.0000
6.250 0.7143 0.02532 0.01392 -0.0005 0.1929 1.0000
6.500 0.7254 0.02696 0.01514 0.0015 0.1533 1.0000
6.750 0.7387 0.02840 0.01645 0.0033 0.1353 1.0000
7.000 0.7527 0.02976 0.01774 0.0051 0.1238 1.0000
7.250 0.7665 0.03112 0.01901 0.0068 0.1152 1.0000
7.500 0.7834 0.03233 0.02034 0.0083 0.1081 1.0000
7.750 0.7992 0.03371 0.02166 0.0098 0.1026 1.0000
8.000 0.8183 0.03498 0.02309 0.0110 0.0969 1.0000
8.250 0.8373 0.03635 0.02449 0.0122 0.0920 1.0000
8.500 0.8578 0.03789 0.02603 0.0130 0.0878 1.0000
8.750 0.8782 0.03945 0.02783 0.0140 0.0837 1.0000
9.000 0.8986 0.04113 0.02962 0.0148 0.0804 1.0000
9.250 0.9201 0.04300 0.03143 0.0153 0.0772 1.0000
9.500 0.9346 0.04505 0.03387 0.0166 0.0745 1.0000
9.750 0.9490 0.04731 0.03644 0.0178 0.0723 1.0000
10.000 0.9621 0.04975 0.03914 0.0190 0.0707 1.0000
10.250 0.9733 0.05219 0.04180 0.0202 0.0691 1.0000
10.500 0.9849 0.05460 0.04431 0.0212 0.0674 1.0000
10.750 0.9884 0.05748 0.04742 0.0226 0.0661 1.0000
11.000 0.9793 0.06070 0.05105 0.0247 0.0652 1.0000
11.250 0.9650 0.06398 0.05464 0.0269 0.0646 1.0000
11.500 0.9474 0.06774 0.05868 0.0282 0.0643 1.0000
11.750 0.9267 0.07221 0.06340 0.0281 0.0641 1.0000
12.000 0.9021 0.07769 0.06911 0.0263 0.0642 1.0000
12.250 0.8737 0.08452 0.07614 0.0228 0.0645 1.0000
12.500 0.8421 0.09303 0.08480 0.0174 0.0650 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NPL 9626 AIRFOIL (npl9626-il)