NLR-1T AIRFOIL (nlr1t-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NLR-1T AIRFOIL (nlr1t-il) Reynolds number: 500,000 Max Cl/Cd: 70.45 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-nlr1t-il-500000.txt Download as CSV file: xf-nlr1t-il-500000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5413 0.08025 0.07807 -0.0327 1.0000 0.0172
-8.750 -0.5482 0.07491 0.07268 -0.0364 1.0000 0.0172
-8.500 -0.5554 0.07045 0.06814 -0.0374 1.0000 0.0172
-8.250 -0.5833 0.06644 0.06411 -0.0337 1.0000 0.0174
-8.000 -0.5843 0.06092 0.05851 -0.0357 0.9964 0.0178
-7.750 -0.5718 0.05722 0.05471 -0.0377 0.9938 0.0181
-7.500 -0.5583 0.05390 0.05129 -0.0391 0.9904 0.0185
-7.250 -0.5423 0.05035 0.04759 -0.0405 0.9875 0.0191
-7.000 -0.5239 0.04654 0.04359 -0.0419 0.9853 0.0200
-6.750 -0.5023 0.04234 0.03911 -0.0429 0.9835 0.0219
-6.500 -0.4784 0.04073 0.03699 -0.0416 0.9799 0.0233
-6.250 -0.4732 0.03358 0.02954 -0.0409 0.9759 0.0245
-6.000 -0.4509 0.03163 0.02752 -0.0413 0.9739 0.0254
-5.750 -0.4274 0.02980 0.02554 -0.0414 0.9722 0.0267
-5.500 -0.4009 0.02846 0.02394 -0.0412 0.9707 0.0299
-5.250 -0.3837 0.02569 0.02066 -0.0393 0.9679 0.0329
-5.000 -0.3624 0.02384 0.01880 -0.0387 0.9655 0.0342
-4.750 -0.3392 0.02254 0.01739 -0.0382 0.9634 0.0360
-4.500 -0.3140 0.02150 0.01617 -0.0375 0.9616 0.0391
-4.250 -0.2846 0.01717 0.01111 -0.0351 0.9601 0.0229
-4.000 -0.2573 0.01536 0.00913 -0.0348 0.9588 0.0224
-3.750 -0.2296 0.01426 0.00793 -0.0347 0.9577 0.0223
-3.500 -0.2014 0.01376 0.00737 -0.0347 0.9567 0.0230
-3.250 -0.1739 0.01313 0.00670 -0.0346 0.9557 0.0233
-3.000 -0.1508 0.01245 0.00598 -0.0336 0.9536 0.0233
-2.750 -0.1294 0.01168 0.00518 -0.0322 0.9503 0.0238
-2.500 -0.1054 0.01104 0.00449 -0.0313 0.9468 0.0252
-2.250 -0.0786 0.01066 0.00410 -0.0309 0.9436 0.0271
-2.000 -0.0532 0.01039 0.00379 -0.0300 0.9385 0.0297
-1.750 -0.0309 0.01002 0.00338 -0.0282 0.9291 0.0359
-1.500 -0.0174 0.00875 0.00294 -0.0249 0.9175 0.2724
-1.250 -0.0112 0.00715 0.00269 -0.0204 0.9072 0.6488
-1.000 0.0039 0.00639 0.00257 -0.0168 0.8993 0.8222
-0.750 0.0934 0.00699 0.00343 -0.0285 0.8973 0.9741
-0.500 0.1536 0.00716 0.00354 -0.0352 0.8932 0.9893
-0.250 0.1919 0.00712 0.00344 -0.0374 0.8876 0.9947
0.000 0.2341 0.00701 0.00328 -0.0405 0.8817 0.9990
0.250 0.2625 0.00691 0.00314 -0.0407 0.8736 1.0000
0.500 0.2869 0.00681 0.00301 -0.0399 0.8649 1.0000
0.750 0.3112 0.00671 0.00285 -0.0390 0.8541 1.0000
1.000 0.3360 0.00661 0.00273 -0.0382 0.8406 1.0000
1.250 0.3611 0.00653 0.00261 -0.0376 0.8242 1.0000
1.500 0.3862 0.00647 0.00250 -0.0369 0.7985 1.0000
1.750 0.4084 0.00669 0.00220 -0.0353 0.6667 1.0000
2.000 0.4317 0.00752 0.00234 -0.0347 0.5156 1.0000
2.250 0.4555 0.00818 0.00249 -0.0343 0.4021 1.0000
2.500 0.4794 0.00873 0.00263 -0.0339 0.3127 1.0000
2.750 0.5036 0.00916 0.00277 -0.0334 0.2544 1.0000
3.000 0.5283 0.00947 0.00292 -0.0330 0.2249 1.0000
3.250 0.5530 0.00972 0.00307 -0.0325 0.2071 1.0000
3.500 0.5778 0.00993 0.00322 -0.0320 0.1953 1.0000
3.750 0.6024 0.01016 0.00340 -0.0315 0.1861 1.0000
4.000 0.6273 0.01032 0.00354 -0.0310 0.1785 1.0000
4.250 0.6515 0.01055 0.00374 -0.0304 0.1717 1.0000
4.500 0.6762 0.01069 0.00390 -0.0298 0.1663 1.0000
4.750 0.7003 0.01092 0.00412 -0.0292 0.1604 1.0000
5.000 0.7247 0.01110 0.00431 -0.0286 0.1550 1.0000
5.250 0.7490 0.01127 0.00449 -0.0280 0.1485 1.0000
5.500 0.7729 0.01149 0.00471 -0.0273 0.1423 1.0000
5.750 0.7973 0.01164 0.00490 -0.0267 0.1353 1.0000
6.000 0.8213 0.01183 0.00510 -0.0261 0.1268 1.0000
6.250 0.8454 0.01200 0.00526 -0.0255 0.1129 1.0000
6.500 0.8676 0.01243 0.00549 -0.0246 0.0799 1.0000
6.750 0.8887 0.01301 0.00599 -0.0234 0.0634 1.0000
7.000 0.9096 0.01359 0.00653 -0.0223 0.0560 1.0000
7.250 0.9312 0.01404 0.00704 -0.0212 0.0528 1.0000
7.500 0.9527 0.01449 0.00755 -0.0200 0.0503 1.0000
7.750 0.9733 0.01502 0.00812 -0.0188 0.0480 1.0000
8.000 0.9922 0.01573 0.00888 -0.0173 0.0452 1.0000
8.250 1.0107 0.01646 0.00968 -0.0157 0.0431 1.0000
8.500 1.0323 0.01679 0.01009 -0.0147 0.0416 1.0000
8.750 1.0523 0.01729 0.01066 -0.0134 0.0401 1.0000
9.000 1.0723 0.01776 0.01120 -0.0122 0.0383 1.0000
9.250 1.0913 0.01828 0.01177 -0.0108 0.0363 1.0000
9.500 1.1055 0.01937 0.01293 -0.0087 0.0342 1.0000
9.750 1.1245 0.01984 0.01349 -0.0074 0.0329 1.0000
10.000 1.1460 0.02001 0.01375 -0.0064 0.0314 1.0000
10.250 1.1662 0.02029 0.01409 -0.0053 0.0295 1.0000
10.500 1.1861 0.02057 0.01442 -0.0041 0.0275 1.0000
10.750 1.2006 0.02136 0.01527 -0.0022 0.0251 1.0000
11.000 1.2263 0.02113 0.01510 -0.0020 0.0220 1.0000
11.250 1.2463 0.02148 0.01542 -0.0009 0.0169 1.0000
11.500 1.2608 0.02229 0.01624 0.0010 0.0133 1.0000
11.750 1.2736 0.02316 0.01718 0.0032 0.0116 1.0000
12.000 1.2818 0.02437 0.01847 0.0060 0.0104 1.0000
12.250 1.2896 0.02550 0.01975 0.0088 0.0098 1.0000
12.500 1.2953 0.02655 0.02095 0.0120 0.0093 1.0000
12.750 1.2998 0.02768 0.02221 0.0151 0.0089 1.0000
13.000 1.3039 0.02893 0.02359 0.0179 0.0085 1.0000
13.250 1.3061 0.03041 0.02519 0.0206 0.0081 1.0000
13.500 1.3049 0.03222 0.02713 0.0232 0.0079 1.0000
13.750 1.2996 0.03444 0.02951 0.0258 0.0076 1.0000
14.000 1.2908 0.03713 0.03237 0.0280 0.0075 1.0000
14.250 1.2794 0.04032 0.03573 0.0296 0.0074 1.0000
14.500 1.2657 0.04411 0.03971 0.0303 0.0073 1.0000
14.750 1.2552 0.04813 0.04390 0.0298 0.0073 1.0000
15.000 1.2420 0.05327 0.04923 0.0277 0.0072 1.0000
15.250 1.2254 0.05998 0.05614 0.0237 0.0072 1.0000
15.500 1.2050 0.06840 0.06475 0.0182 0.0073 1.0000
15.750 1.1808 0.07771 0.07424 0.0124 0.0073 1.0000
16.000 1.1600 0.08616 0.08283 0.0078 0.0074 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NLR-1T AIRFOIL (nlr1t-il)