NLR-7223-62 AIRFOIL (nl722362-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NLR-7223-62 AIRFOIL (nl722362-il) Reynolds number: 200,000 Max Cl/Cd: 57.06 at α=2.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-nl722362-il-200000.txt Download as CSV file: xf-nl722362-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NLR-7223-62 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4887 0.08764 0.08410 -0.0322 1.0000 0.0393
-8.750 -0.5071 0.08183 0.07834 -0.0399 1.0000 0.0401
-8.500 -0.5194 0.07869 0.07521 -0.0395 1.0000 0.0401
-8.250 -0.5432 0.07635 0.07283 -0.0373 1.0000 0.0404
-8.000 -0.5634 0.07459 0.07103 -0.0331 1.0000 0.0405
-7.750 -0.5876 0.07319 0.06949 -0.0282 1.0000 0.0406
-7.500 -0.5999 0.07126 0.06743 -0.0244 1.0000 0.0407
-7.250 -0.6096 0.06603 0.06220 -0.0219 1.0000 0.0413
-7.000 -0.6072 0.06255 0.05883 -0.0198 1.0000 0.0420
-6.750 -0.6060 0.06017 0.05645 -0.0173 1.0000 0.0429
-6.500 -0.6053 0.05774 0.05396 -0.0148 1.0000 0.0437
-6.250 -0.6030 0.05544 0.05157 -0.0123 1.0000 0.0463
-6.000 -0.6023 0.05554 0.05094 -0.0078 1.0000 0.0509
-5.750 -0.6026 0.04932 0.04472 -0.0060 1.0000 0.0523
-5.500 -0.5931 0.04647 0.04198 -0.0045 1.0000 0.0540
-5.250 -0.5834 0.04433 0.03977 -0.0024 1.0000 0.0570
-5.000 -0.5705 0.04626 0.04092 0.0019 1.0000 0.0632
-4.750 -0.5647 0.03925 0.03404 0.0029 1.0000 0.0657
-4.500 -0.5508 0.03707 0.03184 0.0045 1.0000 0.0687
-3.250 -0.4524 0.02485 0.01728 0.0176 1.0000 0.0409
-3.000 -0.4303 0.02309 0.01525 0.0191 1.0000 0.0410
-2.750 -0.4067 0.02091 0.01296 0.0202 1.0000 0.0395
-2.500 -0.3829 0.01937 0.01125 0.0215 1.0000 0.0385
-2.250 -0.3555 0.01829 0.01006 0.0218 0.9990 0.0380
-2.000 -0.3218 0.01744 0.00913 0.0209 0.9962 0.0389
-1.750 -0.2888 0.01686 0.00852 0.0200 0.9929 0.0403
-1.500 -0.2578 0.01613 0.00787 0.0192 0.9890 0.0466
-1.250 -0.2214 0.01577 0.00748 0.0174 0.9853 0.0535
-1.000 -0.0804 0.01303 0.00813 -0.0050 0.9989 1.0000
-0.750 -0.0212 0.01320 0.00812 -0.0118 0.9909 1.0000
-0.500 0.1013 0.01302 0.00774 -0.0308 0.9718 1.0000
-0.250 0.1927 0.01248 0.00711 -0.0433 0.9597 1.0000
0.000 0.2285 0.01230 0.00689 -0.0449 0.9515 1.0000
0.250 0.2882 0.01194 0.00652 -0.0513 0.9468 1.0000
0.500 0.3325 0.01165 0.00622 -0.0545 0.9395 1.0000
0.750 0.3807 0.01122 0.00580 -0.0583 0.9311 1.0000
1.000 0.4124 0.01095 0.00555 -0.0586 0.9194 1.0000
1.250 0.4411 0.01069 0.00531 -0.0582 0.9061 1.0000
1.500 0.4651 0.01047 0.00509 -0.0567 0.8906 1.0000
1.750 0.4865 0.01024 0.00486 -0.0546 0.8725 1.0000
2.000 0.5050 0.01002 0.00465 -0.0520 0.8488 1.0000
2.250 0.5235 0.00978 0.00441 -0.0492 0.8146 1.0000
2.500 0.5404 0.00947 0.00379 -0.0456 0.6903 1.0000
2.750 0.5472 0.01056 0.00367 -0.0406 0.4786 1.0000
3.000 0.5586 0.01161 0.00399 -0.0377 0.3508 1.0000
3.250 0.5759 0.01228 0.00431 -0.0359 0.2956 1.0000
3.500 0.5957 0.01279 0.00464 -0.0344 0.2699 1.0000
3.750 0.6165 0.01321 0.00496 -0.0331 0.2533 1.0000
4.000 0.6375 0.01364 0.00528 -0.0319 0.2412 1.0000
4.250 0.6594 0.01400 0.00561 -0.0308 0.2312 1.0000
4.500 0.6812 0.01440 0.00601 -0.0297 0.2226 1.0000
4.750 0.7031 0.01481 0.00638 -0.0286 0.2148 1.0000
5.000 0.7254 0.01521 0.00681 -0.0276 0.2075 1.0000
5.250 0.7473 0.01561 0.00720 -0.0265 0.1999 1.0000
5.500 0.7695 0.01603 0.00769 -0.0255 0.1925 1.0000
5.750 0.7914 0.01644 0.00810 -0.0244 0.1850 1.0000
6.000 0.8128 0.01681 0.00855 -0.0233 0.1758 1.0000
6.250 0.8331 0.01716 0.00890 -0.0221 0.1650 1.0000
6.500 0.8530 0.01740 0.00916 -0.0207 0.1537 1.0000
6.750 0.8734 0.01758 0.00948 -0.0194 0.1427 1.0000
7.000 0.8933 0.01782 0.00981 -0.0180 0.1311 1.0000
7.250 0.9132 0.01798 0.01006 -0.0166 0.1172 1.0000
7.500 0.9339 0.01804 0.01016 -0.0154 0.0943 1.0000
7.750 0.9487 0.01904 0.01090 -0.0132 0.0661 1.0000
8.000 0.9626 0.02024 0.01213 -0.0108 0.0536 1.0000
8.250 0.9755 0.02145 0.01331 -0.0083 0.0467 1.0000
8.500 0.9917 0.02237 0.01436 -0.0063 0.0423 1.0000
8.750 1.0056 0.02349 0.01550 -0.0040 0.0395 1.0000
9.000 1.0184 0.02522 0.01726 -0.0017 0.0375 1.0000
9.250 1.0344 0.02647 0.01870 0.0003 0.0360 1.0000
9.500 1.0495 0.02785 0.02024 0.0024 0.0345 1.0000
9.750 1.0635 0.02912 0.02161 0.0044 0.0329 1.0000
10.000 1.0770 0.03073 0.02326 0.0063 0.0316 1.0000
10.250 1.0885 0.03351 0.02622 0.0082 0.0306 1.0000
10.500 1.0970 0.03533 0.02833 0.0110 0.0302 1.0000
10.750 1.1016 0.03736 0.03069 0.0142 0.0299 1.0000
11.000 1.1027 0.03962 0.03326 0.0177 0.0297 1.0000
11.250 1.0991 0.04201 0.03596 0.0215 0.0295 1.0000
11.500 1.0889 0.04430 0.03852 0.0263 0.0294 1.0000
11.750 1.0749 0.04675 0.04122 0.0309 0.0295 1.0000
12.000 1.0594 0.04945 0.04417 0.0348 0.0295 1.0000
12.250 1.0434 0.05256 0.04750 0.0377 0.0297 1.0000
12.500 1.0223 0.05615 0.05133 0.0399 0.0297 1.0000
12.750 1.0012 0.06020 0.05558 0.0408 0.0298 1.0000
13.000 0.9817 0.06479 0.06033 0.0404 0.0301 1.0000
13.250 0.9561 0.07068 0.06641 0.0379 0.0301 1.0000
13.500 0.9315 0.07781 0.07368 0.0338 0.0304 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NLR-7223-62 AIRFOIL (nl722362-il)