15x12E (15x12E.dat) v2022-0915 Simulation Date: 11/02/2023 ----- AIRFOIL SUMMARY DATA ----- DEFINITIONS: THE QUOTED PITCH REFLECTS, IN GENERAL, ANGULAR MEASURE AS DEFINED WITH A FLAT BOTTOM SURFACE. THIS WILL AGREE WITH A PRATHER GAGE MEASUREMENT OVER MOST OF THE EFFECTIVE PORTION OF THE BLADE. (NOTE: QUOTED=INPUT) THE LE-TE MEASURE IS DEFINED IN TERMS OF LEADING EDGE AND TRAILING EDGE (MOLD) PARTING LINE DATUMS. THE PRATHER MEASURE REFLECTS THE MOST LIKELY PITCH INTERPRETATION FROM A PITCH MEASUREMENT DEVICE. THAT RESTS AGAINST THE LOWER SURFACE. SWEEP IS DEFINED WITH (MOLD) LE PARTING LINE. ZHIGH IS HIGHEST ELEVATION ON TOP SURFACE. TWIST IS DEFINED WITH (MOLD) LE AND TE PARTING LINE DATUMS. CHORD IS THE LENGTH BETWEEN (MOLD) LE AND TE PARTING LINES. CGY IS MASS OFFSET, FORE-AFT. CGZ IS MASS OFFSET, ELEVATION. PITCH (REAL) IS FORMED WITH CHORD PERPENDICULAR TO NOSE TANGENT STATION CHORD PITCH PITCH PITCH SWEEP THICKNESS TWIST MAX-THICK CROSS-SECTION ZHIGH CGY CGZ PITCH (IN) (IN) (QUOTED) (LE-TE) (PRATHER) (IN) RATIO (DEG) (IN) (IN**2) (IN) (IN) (IN) (REAL) 1.2532 0.9416 10.5846 10.5846 8.8957 0.4635 0.2409 53.3545 0.2268 0.1300 0.6069 0.1959 0.3113 10.3604 1.3434 0.9714 10.9013 10.9013 9.3054 0.4738 0.2294 52.2489 0.2228 0.1316 0.6175 0.1926 0.3189 10.7119 1.4337 0.9999 11.1780 11.1780 9.6779 0.4833 0.2186 51.1351 0.2186 0.1328 0.6263 0.1885 0.3253 11.0189 1.5240 1.0271 11.4146 11.4146 10.0118 0.4920 0.2086 50.0077 0.2143 0.1337 0.6331 0.1835 0.3304 11.2788 1.6142 1.0531 11.6111 11.6111 10.3041 0.4999 0.1994 48.8622 0.2099 0.1344 0.6381 0.1778 0.3343 11.4923 1.7045 1.0779 11.7676 11.7676 10.5532 0.5070 0.1908 47.6947 0.2057 0.1349 0.6412 0.1712 0.3369 11.6659 1.7948 1.1017 11.8840 11.8840 10.7576 0.5133 0.1829 46.5017 0.2015 0.1354 0.6425 0.1637 0.3382 11.7980 1.8970 1.1271 11.9675 11.9675 10.9339 0.5196 0.1747 45.1163 0.1970 0.1359 0.6416 0.1544 0.3381 11.8962 2.0740 1.1650 12.0000 12.0000 11.1079 0.5282 0.1624 42.6411 0.1892 0.1367 0.6344 0.1365 0.3345 11.9491 2.2530 1.1956 12.0000 12.0000 11.2187 0.5342 0.1521 40.2873 0.1818 0.1368 0.6213 0.1189 0.3268 11.9649 2.4321 1.2188 12.0000 12.0000 11.3070 0.5376 0.1436 38.1414 0.1750 0.1365 0.6075 0.1020 0.3193 11.9783 2.6112 1.2350 12.0000 12.0000 11.3768 0.5384 0.1368 36.1822 0.1689 0.1358 0.5936 0.0858 0.3123 11.9924 2.7903 1.2445 12.0000 12.0000 11.4308 0.5368 0.1313 34.3906 0.1634 0.1346 0.5794 0.0704 0.3057 12.0180 2.9693 1.2478 12.0000 12.0000 11.4714 0.5330 0.1270 32.7489 0.1585 0.1332 0.5651 0.0559 0.2994 12.0489 3.1484 1.2452 12.0000 12.0000 11.5016 0.5270 0.1237 31.2415 0.1540 0.1313 0.5507 0.0423 0.2935 12.0768 3.3275 1.2370 12.0000 12.0000 11.5243 0.5189 0.1210 29.8543 0.1496 0.1288 0.5361 0.0295 0.2878 12.0992 3.5066 1.2237 12.0000 12.0000 11.5430 0.5089 0.1186 28.5751 0.1452 0.1257 0.5214 0.0176 0.2823 12.1179 3.6856 1.2057 12.0000 12.0000 11.5599 0.4972 0.1164 27.3928 0.1404 0.1217 0.5066 0.0066 0.2771 12.1340 3.8647 1.1833 12.0000 12.0000 11.5707 0.4837 0.1144 26.2977 0.1354 0.1169 0.4917 -0.0036 0.2720 12.1479 4.0438 1.1569 12.0000 12.0000 11.5675 0.4687 0.1125 25.2812 0.1301 0.1114 0.4766 -0.0130 0.2669 12.1600 4.2228 1.1269 12.0000 12.0000 11.5620 0.4522 0.1107 24.3357 0.1247 0.1053 0.4614 -0.0216 0.2619 12.1708 4.4019 1.0937 12.0000 12.0000 11.5559 0.4344 0.1090 23.4546 0.1192 0.0988 0.4460 -0.0294 0.2570 12.1804 4.5810 1.0576 12.0000 12.0000 11.5500 0.4154 0.1074 22.6317 0.1136 0.0920 0.4306 -0.0365 0.2521 12.1894 4.7601 1.0191 12.0000 12.0000 11.5444 0.3953 0.1060 21.8619 0.1080 0.0849 0.4151 -0.0430 0.2471 12.1983 4.9391 0.9785 12.0000 12.0000 11.5393 0.3743 0.1047 21.1404 0.1025 0.0778 0.3994 -0.0490 0.2421 12.2080 5.1182 0.9361 12.0000 12.0000 11.5346 0.3524 0.1035 20.4630 0.0969 0.0707 0.3836 -0.0544 0.2369 12.2190 5.2973 0.8925 12.0000 12.0000 11.5302 0.3298 0.1025 19.8260 0.0915 0.0638 0.3677 -0.0596 0.2314 12.2344 5.4764 0.8479 12.0000 12.0000 11.5253 0.3066 0.1016 19.2260 0.0861 0.0573 0.3517 -0.0648 0.2254 12.2487 5.6554 0.8028 12.0000 12.0000 11.5172 0.2829 0.1008 18.6600 0.0809 0.0510 0.3355 -0.0699 0.2187 12.2651 5.8345 0.7574 12.0000 12.0000 11.5079 0.2588 0.1001 18.1253 0.0758 0.0452 0.3192 -0.0750 0.2116 12.2825 6.0136 0.7123 12.0000 12.0000 11.4980 0.2345 0.0996 17.6194 0.0709 0.0399 0.3029 -0.0802 0.2039 12.3014 6.1926 0.6677 12.0000 12.0000 11.4872 0.2101 0.0992 17.1402 0.0662 0.0349 0.2864 -0.0857 0.1956 12.3219 6.3717 0.6242 12.0000 12.0000 11.4756 0.1856 0.0989 16.6856 0.0617 0.0305 0.2699 -0.0914 0.1867 12.3485 6.5508 0.5819 12.0000 12.0000 11.4630 0.1613 0.0987 16.2538 0.0574 0.0264 0.2532 -0.0975 0.1772 12.3772 6.7299 0.5414 12.0000 12.0000 11.4482 0.1372 0.0987 15.8433 0.0534 0.0229 0.2365 -0.1040 0.1671 12.4074 6.9089 0.5030 12.0000 12.0000 11.4313 0.1135 0.0988 15.4526 0.0497 0.0197 0.2197 -0.1109 0.1564 12.4387 7.0880 0.4670 12.0000 12.0000 11.3948 0.0902 0.0990 15.0802 0.0462 0.0170 0.2025 -0.1183 0.1447 12.4505 7.2597 0.4273 12.0000 12.0000 11.2233 0.0607 0.0993 14.7392 0.0424 0.0143 0.1819 -0.1302 0.1290 12.2974 7.4250 0.2747 12.0000 12.0000 10.1090 -0.0805 0.0998 14.4249 0.0274 0.0059 0.1287 -0.2037 0.0937 11.9374 7.5000 0.0139 12.0000 12.0301 12.0301 -0.3310 0.1000 14.3209 0.0014 0.0000 0.0541 0.0000 0.0000 0.0000 RADIUS: 7.50 PROPELLER RADIUS (IN) HUBTRA: 1.25 HUB TRANSITION (IN) BLADES: 2 NUMBER OF BLADES ----- INERTIA AND AREA DATA ----- TOTAL WEIGHT (LB) = 0.099018 TOTAL WEIGHT (Kg) = 0.044906 TOTAL VOLUME (IN**3) = 1.613137 TOTAL PROJECTED AREA (IN**2) = 11.9463 MOMENT OF INERTIA (SNAIL-IN**2) = 0.002494 MOMENT OF INERTIA (Kg-M**2) = 0.000282 STATIC MOMENT, ONE SIDE (IN-LB) = 0.124932 SANITY CHECK DATA ARE BELOW DENSITY (SPECIFIC GRAVITY, INPUT FILE) = 1.700 DENSITY (INPUT FILE, LB/IN**3) = 0.6138E-01 AVERAGE DENSITY (FROM WT & VOL ABOVE, LB/IN**3) = 0.6138E-01 ACTIVITY FACTOR = 73.680 INNER LIMIT (NORMALIZED) = 0.067 ----- NATURAL FREQUENCY DATA ----- LOWEST NATURAL BENDING FREQUENCY (IN TERMS OF RPM) = 8267.20 BASED ON MODULUS (MILLION) = 2.70 AND, MATERIAL DENSITY (S.G.) = 1.70 ----- AIRFOIL SECTIONS ----- AIRFOIL1: 1.20, E63 (Transition Start, Airfoil 1) AIRFOIL2: 5.12, APC12 (Transition End, Airfoil 2) AIRFOILS ARE SCALED BASED ON THICKNESS RATIO IN TABLE ABOVE. NOTE: APC12 airfoil is equivalent to NACA 4412