NACA 5-H-10 AIRFOIL (n5h10-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NACA 5-H-10 AIRFOIL (n5h10-il) Reynolds number: 500,000 Max Cl/Cd: 70.02 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n5h10-il-500000.txt Download as CSV file: xf-n5h10-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 5-H-10 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.5076 0.08715 0.08462 -0.0204 0.8411 0.0188
-9.000 -0.5156 0.08211 0.07959 -0.0243 0.8356 0.0188
-8.750 -0.5264 0.07765 0.07507 -0.0259 0.8303 0.0188
-8.500 -0.5362 0.07393 0.07130 -0.0255 0.8251 0.0188
-8.250 -0.5549 0.06763 0.06492 -0.0254 0.8202 0.0190
-8.000 -0.5632 0.06318 0.06038 -0.0246 0.8159 0.0194
-7.750 -0.5606 0.06011 0.05724 -0.0239 0.8115 0.0197
-7.500 -0.5538 0.05768 0.05474 -0.0233 0.8068 0.0201
-7.250 -0.5455 0.05522 0.05218 -0.0225 0.8026 0.0207
-7.000 -0.5361 0.05248 0.04931 -0.0218 0.7985 0.0217
-6.750 -0.5265 0.04884 0.04551 -0.0208 0.7937 0.0228
-6.500 -0.5067 0.04597 0.04226 -0.0190 0.7894 0.0254
-6.250 -0.4955 0.04266 0.03864 -0.0170 0.7855 0.0255
-6.000 -0.4841 0.03901 0.03469 -0.0150 0.7806 0.0256
-5.750 -0.4709 0.03556 0.03092 -0.0130 0.7762 0.0256
-5.500 -0.4803 0.02150 0.01558 -0.0067 0.7733 0.0198
-5.250 -0.4549 0.02199 0.01619 -0.0068 0.7686 0.0213
-5.000 -0.4312 0.02061 0.01459 -0.0060 0.7637 0.0225
-4.750 -0.4070 0.01882 0.01248 -0.0050 0.7597 0.0236
-4.500 -0.3811 0.01781 0.01125 -0.0043 0.7551 0.0249
-4.250 -0.3555 0.01597 0.00917 -0.0038 0.7501 0.0268
-4.000 -0.3298 0.01522 0.00837 -0.0035 0.7459 0.0290
-3.750 -0.3033 0.01454 0.00760 -0.0032 0.7417 0.0308
-3.500 -0.2765 0.01386 0.00686 -0.0029 0.7365 0.0327
-3.250 -0.2496 0.01361 0.00653 -0.0026 0.7321 0.0347
-3.000 -0.2249 0.01243 0.00524 -0.0020 0.7281 0.0370
-2.750 -0.1997 0.01181 0.00462 -0.0015 0.7229 0.0392
-2.500 -0.1740 0.01145 0.00423 -0.0011 0.7183 0.0418
-2.250 -0.1483 0.01114 0.00386 -0.0006 0.7145 0.0441
-2.000 -0.1223 0.01082 0.00351 -0.0002 0.7095 0.0457
-1.750 -0.0966 0.01050 0.00315 0.0002 0.7051 0.0475
-1.500 -0.0713 0.01016 0.00278 0.0008 0.7015 0.0529
-1.250 -0.0448 0.00997 0.00258 0.0011 0.6966 0.0580
-1.000 -0.0191 0.00968 0.00234 0.0015 0.6915 0.0766
-0.750 -0.0091 0.00805 0.00206 0.0044 0.6873 0.4675
-0.500 -0.0092 0.00689 0.00225 0.0108 0.6827 0.8581
-0.250 0.0137 0.00707 0.00245 0.0125 0.6780 0.8898
0.000 0.0369 0.00722 0.00258 0.0141 0.6742 0.9111
0.250 0.0616 0.00738 0.00274 0.0155 0.6704 0.9269
0.500 0.0943 0.00762 0.00298 0.0151 0.6661 0.9399
0.750 0.1513 0.00807 0.00338 0.0096 0.6621 0.9485
1.000 0.2344 0.00870 0.00392 -0.0016 0.6586 0.9588
1.250 0.2784 0.00881 0.00402 -0.0050 0.6542 0.9617
1.500 0.3081 0.00878 0.00397 -0.0055 0.6477 0.9634
1.750 0.3361 0.00876 0.00392 -0.0057 0.6401 0.9653
2.000 0.3607 0.00871 0.00383 -0.0051 0.6317 0.9666
2.250 0.3834 0.00864 0.00377 -0.0041 0.6230 0.9672
2.500 0.4057 0.00859 0.00367 -0.0031 0.6144 0.9677
2.750 0.4275 0.00851 0.00362 -0.0020 0.6050 0.9684
3.000 0.4480 0.00848 0.00358 -0.0006 0.5978 0.9693
3.250 0.4724 0.00844 0.00356 0.0000 0.5892 0.9701
3.500 0.4992 0.00839 0.00352 0.0001 0.5761 0.9705
3.750 0.5258 0.00837 0.00349 0.0001 0.5599 0.9708
4.000 0.5524 0.00837 0.00348 0.0002 0.5417 0.9713
4.250 0.5787 0.00841 0.00350 0.0003 0.5145 0.9718
4.500 0.6036 0.00862 0.00354 0.0006 0.4560 0.9724
4.750 0.6221 0.00967 0.00397 0.0014 0.3194 0.9734
5.000 0.6390 0.01100 0.00458 0.0021 0.1632 0.9747
5.250 0.6602 0.01176 0.00507 0.0026 0.0997 0.9759
5.500 0.6823 0.01245 0.00550 0.0030 0.0519 0.9774
5.750 0.7056 0.01288 0.00589 0.0034 0.0428 0.9789
6.000 0.7278 0.01325 0.00633 0.0040 0.0398 0.9805
6.250 0.7468 0.01373 0.00684 0.0053 0.0368 0.9824
6.500 0.7691 0.01457 0.00779 0.0055 0.0328 0.9838
6.750 0.7957 0.01509 0.00838 0.0051 0.0314 0.9847
7.000 0.8210 0.01575 0.00911 0.0048 0.0291 0.9859
7.250 0.8448 0.01652 0.00990 0.0046 0.0262 0.9874
7.500 0.8598 0.01845 0.01193 0.0055 0.0237 0.9900
7.750 0.8847 0.01893 0.01248 0.0054 0.0227 0.9918
8.000 0.9077 0.01963 0.01326 0.0056 0.0214 0.9937
8.250 0.9325 0.02048 0.01417 0.0053 0.0200 0.9953
8.500 0.9586 0.02115 0.01486 0.0045 0.0184 0.9970
8.750 0.9781 0.02364 0.01739 0.0049 0.0169 0.9983
9.000 1.0041 0.02457 0.01846 0.0044 0.0164 0.9998
9.250 1.0078 0.02502 0.01902 0.0082 0.0158 1.0000
9.500 1.0057 0.02585 0.01994 0.0132 0.0156 1.0000
9.750 1.0145 0.02689 0.02110 0.0160 0.0150 1.0000
10.000 1.0268 0.02833 0.02268 0.0180 0.0145 1.0000
10.250 1.0379 0.02952 0.02398 0.0200 0.0139 1.0000
10.500 1.0448 0.03011 0.02461 0.0226 0.0133 1.0000
10.750 1.0505 0.03119 0.02580 0.0250 0.0130 1.0000
11.000 1.0559 0.03271 0.02736 0.0270 0.0123 1.0000
11.250 1.0530 0.03701 0.03202 0.0294 0.0120 1.0000
11.500 1.0497 0.03931 0.03455 0.0317 0.0119 1.0000
11.750 1.0447 0.04159 0.03706 0.0338 0.0117 1.0000
12.000 1.0375 0.04440 0.04009 0.0355 0.0117 1.0000
12.250 1.0258 0.04789 0.04381 0.0369 0.0117 1.0000
12.500 1.0116 0.05153 0.04769 0.0378 0.0116 1.0000
12.750 0.9972 0.05532 0.05169 0.0382 0.0114 1.0000
13.000 0.9787 0.05998 0.05654 0.0379 0.0116 1.0000
13.250 0.9607 0.06476 0.06151 0.0368 0.0114 1.0000
13.500 0.9379 0.07068 0.06761 0.0349 0.0115 1.0000
13.750 0.9181 0.07658 0.07364 0.0324 0.0117 1.0000
14.000 0.8131 0.07586 0.07305 0.0346 0.0121 1.0000
14.250 0.7972 0.08201 0.07933 0.0317 0.0122 1.0000
14.500 0.7772 0.08929 0.08675 0.0278 0.0123 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 5-H-10 AIRFOIL (n5h10-il)