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• Aerodynamic Characteristics

Piano calculates the complete aerodynamic Lift-Drag characteristics (the 'polar') of an aircraft from its geometric description and allowing for various technology-level parameters. It is possible to modify the shape of the polar through factors and/or tabulated adjustments. Alternatively, the user can override all internal calculations and explicitly input aerodynamic lookup tables in a variety of common formats. Piano can then be used to analyse or verify the performance of known, flight-tested aircraft.

Detailed classical drag-buildup techniques are implemented, as applicable to preliminary design. The zero-lift drag calculations account for all wetted areas, skin friction coefficients and empirical form factors. The wing is split spanwise into panels and Reynolds effects are included under all conditions. Fuselage, nacelle external (scrubbing) drag, and vertical / horizontal tail drag are similarly calculated. Lift-dependent drag has been calibrated using known polars. It is corrected for trim requirements and can include simple adjustments for winglets.

High-speed compressibility drag and divergence Mach are derived from methods originally developed at the RAE and in industry, accounting for different levels of supercritical or conventional aerofoil technology. Additional adjustments for non-lifting component compressibility contributions and drag 'creep' are available. Statistical buffet boundaries are estimated (or can be input).

Low-speed aerodynamics are evaluated through a mixture of textbook methods, with a choice of commonly-used flap types. Factors on the estimated overall CLmax and low-speed L/D ratios are routinely used to adjust these values which are sensitive to configuration details and may require wind tunnel or flight-test verification.

The following sample breakdown shows the basic drag contributions:

 MACH                  0.800
 Altitude (pressure)   45000. feet
 KTAS                  458.9
 KEAS                  201.9
 KCAS                  215.5
 Reynolds number       1.200  millions per foot
 Delta-ISA               +0.  deg.C.

 CL                    0.478  based on:
 Reference Area       1137.00 sq.feet  (trapezoidal)

 Drag Coefficients based on ref.area

 Cd Zero-Lift              0.01537    (58.6 %)
 Cd Lift-Induced           0.01009    (38.4 %)
 Cd Compressibility        0.00059    ( 2.3 %)
 Cd Trim                   0.00019    ( 0.7 %)
 Delta Cd (polar-mod)      0.00000    ( 0.0 %)
                           -------    --------
 Cd Total                  0.02625    ( 100 %)

 Aerodynamic Boundaries:
 Divergence Mach         0.803 {at the given CL 0.478}
 Initial Buffet Mach     0.879 {at the given CL 0.478}
 Initial Buffet CL       0.775 {at the given Mach 0.800}

 Zero-Lift Component Breakdown (Drag Areas,= Cd*S = D/q)

 Wing                    6.398  sq.feet    (36.6 %)
 Winglets                0.203  sq.feet    ( 1.2 %)
 Fuselage & fairing      4.949  sq.feet    (28.3 %)
 Stabiliser              2.257  sq.feet    (12.9 %)
 Fin                     1.206  sq.feet    ( 6.9 %)  (incl.dorsal)
 Nacelles (total)        2.466  sq.feet    (14.1 %)
 User CdS Increment      0.000  sq.feet    ( 0.0 %)
                       -------             --------
 Total Cd0*S            17.479  sq.feet    ( 100 %)

 Overall Lift / Drag Ratio = 18.21 

 Total Lift Force       75000. lbf.
 Total Drag Force        4118. lbf. (2059.lbf. per engine) 

 Engine / Airframe Performance:
 Total Fuel Flow at Thrust=Drag:       2638. lb/hr
 Specific Fuel Consumption (SFC)       0.6406 lb/hr/lbf
 Specific Air Range (SAR)              0.1739 nm/lb
 Available Total Thrust at MCR:        4693. lbf.
 Available Total Thrust at MCL:        5130. lbf.

 Instantaneous Performance Reserves at Weight=Lift:
 RoC at MCR, constant Mach 0.800:      356. feet/min
 RoC at MCR, constant 216.KCAS:        257. feet/min

 RoC at MCL, constant Mach 0.800:      627. feet/min
 RoC at MCL, constant 216.KCAS:        452. feet/min

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