5x7.5E                            (5x75E.dat)               
v2022-0215
Simulation Date: 05/18/2022


       ----- 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.
         

      STATION     CHORD       PITCH       PITCH        PITCH       SWEEP    THICKNESS      TWIST      MAX-THICK  CROSS-SECTION ZHIGH       CGY          CGZ                        
       (IN)       (IN)       (QUOTED)    (LE-TE)     (PRATHER)      (IN)     RATIO         (DEG)       (IN)      (IN**2)      (IN)         (IN)         (IN)                 

      0.6630      0.5499      6.3381      6.3381      5.6848      0.2049      0.1249     56.6834      0.0687      0.0315      0.2723      0.0599      0.0803
      0.6932      0.5485      6.6828      6.6828      6.1364      0.2061      0.1232     56.9056      0.0676      0.0295      0.2702      0.0617      0.0803
      0.7234      0.5470      6.9669      6.9669      6.5227      0.2072      0.1216     56.8808      0.0665      0.0278      0.2680      0.0626      0.0803
      0.7535      0.5456      7.1906      7.1906      6.8401      0.2083      0.1201     56.6379      0.0655      0.0263      0.2656      0.0627      0.0802
      0.7837      0.5442      7.3538      7.3538      7.0856      0.2092      0.1186     56.1942      0.0645      0.0250      0.2631      0.0620      0.0800
      0.8138      0.5428      7.4565      7.4565      7.2571      0.2101      0.1172     55.5586      0.0636      0.0240      0.2604      0.0607      0.0797
      0.8490      0.5411      7.5000      7.5000      7.3629      0.2109      0.1156     54.5788      0.0626      0.0230      0.2570      0.0583      0.0794
      0.9029      0.5384      7.5000      7.5000      7.4191      0.2121      0.1134     52.8956      0.0611      0.0221      0.2514      0.0540      0.0780
      0.9613      0.5354      7.5000      7.5000      7.4318      0.2129      0.1112     51.1540      0.0595      0.0218      0.2447      0.0498      0.0751
      1.0200      0.5321      7.5000      7.5000      7.4267      0.2135      0.1092     49.4871      0.0581      0.0217      0.2373      0.0457      0.0711
      1.0786      0.5285      7.5000      7.5000      7.4198      0.2136      0.1075     47.8988      0.0568      0.0216      0.2297      0.0418      0.0668
      1.1373      0.5247      7.5000      7.5000      7.4082      0.2135      0.1060     46.3864      0.0556      0.0215      0.2221      0.0380      0.0625
      1.1959      0.5204      7.5000      7.5000      7.3969      0.2130      0.1047     44.9463      0.0545      0.0213      0.2144      0.0344      0.0583
      1.2545      0.5157      7.5000      7.5000      7.3870      0.2121      0.1037     43.5753      0.0535      0.0211      0.2065      0.0309      0.0540
      1.3132      0.5105      7.5000      7.5000      7.3786      0.2108      0.1029     42.2700      0.0525      0.0209      0.1987      0.0276      0.0501
      1.3719      0.5047      7.5000      7.5000      7.3722      0.2092      0.1024     41.0269      0.0517      0.0206      0.1909      0.0246      0.0464
      1.4305      0.4983      7.5000      7.5000      7.3679      0.2073      0.1021     39.8429      0.0509      0.0202      0.1831      0.0219      0.0430
      1.4892      0.4913      7.5000      7.5000      7.3656      0.2049      0.1020     38.7147      0.0501      0.0197      0.1752      0.0195      0.0401
      1.5478      0.4835      7.5000      7.5000      7.3647      0.2022      0.1019     37.6394      0.0493      0.0191      0.1673      0.0173      0.0374
      1.6064      0.4750      7.5000      7.5000      7.3640      0.1991      0.1017     36.6139      0.0483      0.0184      0.1593      0.0152      0.0346
      1.6651      0.4657      7.5000      7.5000      7.3635      0.1957      0.1016     35.6356      0.0473      0.0177      0.1513      0.0134      0.0318
      1.7237      0.4554      7.5000      7.5000      7.3632      0.1918      0.1015     34.7018      0.0462      0.0169      0.1432      0.0117      0.0289
      1.7824      0.4443      7.5000      7.5000      7.3632      0.1876      0.1014     33.8100      0.0451      0.0161      0.1351      0.0101      0.0260
      1.8411      0.4321      7.5000      7.5000      7.3637      0.1830      0.1013     32.9577      0.0438      0.0152      0.1270      0.0088      0.0231
      1.8997      0.4189      7.5000      7.5000      7.3648      0.1780      0.1012     32.1429      0.0424      0.0143      0.1188      0.0076      0.0201
      1.9583      0.4046      7.5000      7.5000      7.3666      0.1726      0.1011     31.3633      0.0409      0.0133      0.1106      0.0067      0.0172
      2.0170      0.3891      7.5000      7.5000      7.3692      0.1668      0.1010     30.6171      0.0393      0.0123      0.1023      0.0060      0.0142
      2.0756      0.3724      7.5000      7.5000      7.3728      0.1606      0.1008     29.9023      0.0376      0.0113      0.0940      0.0055      0.0113
      2.1343      0.3544      7.5000      7.5000      7.3779      0.1540      0.1007     29.2172      0.0357      0.0103      0.0857      0.0053      0.0083
      2.1930      0.3351      7.5000      7.5000      7.3848      0.1470      0.1006     28.5602      0.0337      0.0092      0.0773      0.0053      0.0054
      2.2516      0.3145      7.5000      7.5000      7.3938      0.1396      0.1005     27.9298      0.0316      0.0081      0.0688      0.0057      0.0025
      2.3103      0.2923      7.5000      7.5000      7.4060      0.1319      0.1004     27.3245      0.0294      0.0070      0.0603      0.0063     -0.0004
      2.3688      0.2672      7.5000      7.5000      7.4222      0.1222      0.1003     26.7439      0.0268      0.0059      0.0511      0.0065     -0.0036
      2.4209      0.2242      7.5000      7.5000      7.4415      0.0960      0.1002     26.2465      0.0225      0.0042      0.0338     -0.0019     -0.0116
      2.4672      0.1466      7.5000      7.5000      7.5688      0.0378      0.1001     25.8185      0.0147      0.0018      0.0009     -0.0273     -0.0288
      2.5000      0.0001      7.5000      7.5497      7.5483     -0.0859      0.1000     25.6703      0.0000      0.0000     -0.0631      0.0000      0.0000


 RADIUS:  2.50    PROPELLER RADIUS (IN)
 HUBTRA:  0.66    HUB TRANSITION (IN)
 BLADES:  2       NUMBER OF BLADES


       ----- INERTIA AND AREA DATA -----

 TOTAL WEIGHT (LB)          =     0.008984
 TOTAL WEIGHT (Kg)          =     0.004075
 TOTAL VOLUME (IN**3)       =     0.146367
 TOTAL PROJECTED AREA (IN**2) =       1.7249
 MOMENT OF INERTIA (SNAIL-IN**2) =     0.000022
 MOMENT OF INERTIA (Kg-M**2) =     0.000003
 STATIC MOMENT, ONE SIDE (IN-LB) =     0.003344

 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 =    114.953
 INNER LIMIT (NORMALIZED) =      0.100



       ----- NATURAL FREQUENCY DATA -----

 LOWEST NATURAL BENDING FREQUENCY (IN TERMS OF RPM) =     31011.08

 BASED ON MODULUS (MILLION)   =    2.80
 AND, MATERIAL DENSITY (S.G.) =     1.70


       ----- AIRFOIL SECTIONS -----

 AIRFOIL1:  0.68, E63         (Transition Start, Airfoil 1)
 AIRFOIL2:  1.50, APC12       (Transition End, Airfoil 2)

 AIRFOILS ARE SCALED BASED ON THICKNESS RATIO IN TABLE ABOVE.

 NOTE: APC12 airfoil is equivalent to NACA 4412

