NLR-1T AIRFOIL (nlr1t-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NLR-1T AIRFOIL (nlr1t-il) Reynolds number: 500,000 Max Cl/Cd: 68.22 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-nlr1t-il-500000-n5.txt Download as CSV file: xf-nlr1t-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NLR-1T AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.5556 0.09364 0.09139 -0.0149 1.0000 0.0098
-9.500 -0.5567 0.08912 0.08690 -0.0175 1.0000 0.0097
-9.250 -0.5594 0.08397 0.08177 -0.0211 1.0000 0.0096
-9.000 -0.5646 0.07808 0.07591 -0.0266 1.0000 0.0095
-8.750 -0.5738 0.07171 0.06952 -0.0332 0.9907 0.0094
-8.500 -0.5750 0.06496 0.06265 -0.0385 0.9691 0.0093
-8.250 -0.5753 0.05879 0.05631 -0.0411 0.9570 0.0092
-8.000 -0.5748 0.05308 0.05038 -0.0418 0.9480 0.0093
-7.750 -0.5725 0.04747 0.04451 -0.0413 0.9413 0.0093
-7.500 -0.5675 0.04210 0.03883 -0.0400 0.9358 0.0095
-7.250 -0.5598 0.03714 0.03352 -0.0382 0.9318 0.0096
-7.000 -0.5496 0.03235 0.02832 -0.0361 0.9280 0.0098
-6.750 -0.5335 0.02899 0.02453 -0.0342 0.9249 0.0102
-6.500 -0.5184 0.02597 0.02110 -0.0324 0.9220 0.0102
-6.250 -0.5006 0.02346 0.01823 -0.0308 0.9195 0.0102
-6.000 -0.4820 0.02083 0.01520 -0.0293 0.9174 0.0103
-5.750 -0.4606 0.01867 0.01275 -0.0283 0.9154 0.0106
-5.500 -0.4367 0.01734 0.01124 -0.0276 0.9137 0.0108
-5.250 -0.4119 0.01637 0.01014 -0.0272 0.9120 0.0111
-5.000 -0.3865 0.01555 0.00921 -0.0267 0.9105 0.0115
-4.750 -0.3609 0.01479 0.00835 -0.0263 0.9090 0.0118
-4.500 -0.3352 0.01409 0.00757 -0.0259 0.9077 0.0121
-4.250 -0.3096 0.01345 0.00684 -0.0255 0.9064 0.0124
-4.000 -0.2839 0.01290 0.00622 -0.0250 0.9051 0.0128
-3.750 -0.2580 0.01247 0.00574 -0.0246 0.9038 0.0134
-3.500 -0.2321 0.01201 0.00524 -0.0242 0.9025 0.0139
-3.250 -0.2065 0.01152 0.00473 -0.0238 0.9008 0.0141
-3.000 -0.1811 0.01111 0.00427 -0.0233 0.8983 0.0143
-2.750 -0.1555 0.01076 0.00388 -0.0228 0.8954 0.0145
-2.500 -0.1299 0.01039 0.00347 -0.0222 0.8931 0.0150
-2.250 -0.1040 0.01007 0.00311 -0.0218 0.8910 0.0161
-2.000 -0.0776 0.00985 0.00287 -0.0214 0.8882 0.0174
-1.750 -0.0520 0.00965 0.00264 -0.0207 0.8808 0.0188
-1.500 -0.0276 0.00943 0.00236 -0.0196 0.8676 0.0206
-1.250 -0.0037 0.00912 0.00208 -0.0185 0.8501 0.0365
-1.000 0.0198 0.00857 0.00185 -0.0176 0.8373 0.1417
-0.750 0.0380 0.00736 0.00165 -0.0161 0.8304 0.4422
-0.500 0.0498 0.00616 0.00158 -0.0128 0.8233 0.7416
-0.250 0.1148 0.00607 0.00210 -0.0199 0.8139 0.9490
0.250 0.1849 0.00634 0.00223 -0.0226 0.7715 0.9761
0.500 0.2153 0.00690 0.00208 -0.0230 0.6196 0.9808
0.750 0.2438 0.00743 0.00219 -0.0234 0.5260 0.9852
1.000 0.2771 0.00786 0.00226 -0.0250 0.4439 0.9871
1.250 0.3095 0.00823 0.00232 -0.0263 0.3759 0.9892
1.500 0.3411 0.00855 0.00239 -0.0274 0.3208 0.9915
1.750 0.3719 0.00886 0.00247 -0.0284 0.2734 0.9938
2.000 0.4039 0.00911 0.00254 -0.0296 0.2369 0.9953
2.500 0.4686 0.00947 0.00269 -0.0320 0.1973 0.9983
2.750 0.5008 0.00963 0.00278 -0.0332 0.1854 0.9997
3.000 0.5274 0.00978 0.00289 -0.0331 0.1758 1.0000
3.250 0.5525 0.00995 0.00300 -0.0326 0.1662 1.0000
3.500 0.5774 0.01011 0.00312 -0.0322 0.1576 1.0000
3.750 0.6024 0.01026 0.00325 -0.0317 0.1522 1.0000
4.000 0.6273 0.01041 0.00340 -0.0312 0.1475 1.0000
4.250 0.6517 0.01060 0.00355 -0.0306 0.1414 1.0000
4.500 0.6765 0.01074 0.00370 -0.0301 0.1353 1.0000
4.750 0.7009 0.01092 0.00387 -0.0295 0.1299 1.0000
5.000 0.7252 0.01110 0.00406 -0.0289 0.1251 1.0000
5.250 0.7497 0.01125 0.00424 -0.0283 0.1189 1.0000
5.500 0.7738 0.01146 0.00443 -0.0277 0.1110 1.0000
5.750 0.7975 0.01169 0.00461 -0.0270 0.0955 1.0000
6.000 0.8199 0.01210 0.00490 -0.0262 0.0719 1.0000
6.250 0.8420 0.01256 0.00526 -0.0253 0.0556 1.0000
6.500 0.8645 0.01293 0.00562 -0.0244 0.0488 1.0000
6.750 0.8870 0.01329 0.00599 -0.0235 0.0448 1.0000
7.000 0.9091 0.01370 0.00641 -0.0225 0.0412 1.0000
7.250 0.9316 0.01403 0.00681 -0.0216 0.0396 1.0000
7.500 0.9539 0.01435 0.00720 -0.0207 0.0380 1.0000
7.750 0.9758 0.01472 0.00760 -0.0197 0.0358 1.0000
8.000 0.9972 0.01513 0.00803 -0.0187 0.0334 1.0000
8.250 1.0177 0.01563 0.00857 -0.0175 0.0311 1.0000
8.500 1.0386 0.01606 0.00908 -0.0163 0.0300 1.0000
8.750 1.0599 0.01640 0.00949 -0.0153 0.0291 1.0000
9.000 1.0810 0.01676 0.00991 -0.0142 0.0279 1.0000
9.250 1.1017 0.01713 0.01034 -0.0131 0.0263 1.0000
9.500 1.1220 0.01754 0.01077 -0.0120 0.0242 1.0000
9.750 1.1411 0.01805 0.01130 -0.0106 0.0219 1.0000
10.000 1.1613 0.01842 0.01177 -0.0094 0.0206 1.0000
10.250 1.1808 0.01884 0.01225 -0.0082 0.0182 1.0000
10.500 1.1992 0.01936 0.01278 -0.0068 0.0146 1.0000
11.000 1.2290 0.02102 0.01445 -0.0030 0.0069 1.0000
11.250 1.2427 0.02190 0.01542 -0.0009 0.0059 1.0000
11.500 1.2567 0.02270 0.01634 0.0011 0.0055 1.0000
11.750 1.2698 0.02354 0.01731 0.0032 0.0052 1.0000
12.000 1.2817 0.02443 0.01834 0.0055 0.0050 1.0000
12.250 1.2921 0.02540 0.01944 0.0078 0.0047 1.0000
12.500 1.2994 0.02643 0.02061 0.0107 0.0045 1.0000
12.750 1.3042 0.02755 0.02186 0.0137 0.0043 1.0000
13.000 1.3079 0.02884 0.02329 0.0167 0.0041 1.0000
13.250 1.3098 0.03035 0.02495 0.0194 0.0039 1.0000
13.500 1.3093 0.03212 0.02689 0.0221 0.0037 1.0000
13.750 1.3071 0.03410 0.02904 0.0245 0.0036 1.0000
14.000 1.3066 0.03605 0.03114 0.0264 0.0036 1.0000
14.250 1.3038 0.03833 0.03358 0.0280 0.0036 1.0000
14.500 1.2986 0.04101 0.03643 0.0292 0.0035 1.0000
14.750 1.2902 0.04429 0.03988 0.0297 0.0035 1.0000
15.000 1.2786 0.04842 0.04421 0.0291 0.0035 1.0000
15.250 1.2624 0.05410 0.05010 0.0265 0.0035 1.0000
15.500 1.2386 0.06265 0.05886 0.0209 0.0035 1.0000
15.750 1.2013 0.07561 0.07207 0.0122 0.0036 1.0000
16.000 1.1514 0.09000 0.08665 0.0045 0.0036 1.0000
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