الفهرس 
Lifting Surface Theory and Vortex Lattice MethodOnly 14 pages are availabe for public view
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Abstract
Traditionally, early in the design studies, when many
concepts of aircraft configurations are being considered,
designers use their experience and historical data to include
their considerations in the concept. However, this approach is
limited to more conventional configurations and can be very
time consuming for this stage of the design process. Once the
specialists get involved, more detailed numerical
methodologies are used. However, those methods cannot yet
respond to the “dozen a day” type configuration evaluations
desired in the initial conceptual design stages.
The aim of this work is to develop an easy to use and low
time consuming solver for predicting the aerodynamic
characteristics and estimating the stability and control
derivatives of a complete configuration aircraft in subsonic
flow regime. The developed solver is integrated with pre- and
post- processing modules to prepare the geometric data of the
computational model and visualize the results respectively.
This solver can be used easily by designers for fast evaluation
of any aircraft configuration in the initial conceptual design
stages.
The present solver computes the numerical solution for the
integral form of Laplace’s equation using the vortex lattice
method by employing vortex ring elements. The Boit-Savart
law is used to compute the induced velocity and the Kutta-
Jukowski theorem is applied to compute the force at each
vortex ring element. The boundary conditions are that there is
no normal velocity component over the surface and the
aircrafts operate in subsonic flow at low angles of attack flight regime In order to validate the developed potential flow solver,
three different wing planforms were used as the test cases.
First, a rectangular wing of AR = 2, b = 12 m, and CMAC = 6 m
at α = 5 and 7 degrees. Second, a swept back wing of TR = 1,
SA = 450, AR =5, b = 10 m, CMAC = 2 m, and S = 20 m2 at
angles of attack in the range from 0 to 10 degrees. Finally, a
swept tapered wing of TR = 0.5, SA = 450, AR = 2, α = 5.00, b
= 12 m, CMAC = 6 m, and S = 72 m2. All the results were
compared with the published experimental and computational
results.
The influences of varying the different parameters on the
computed force and moment coefficients, and the load
distribution over the wing span were studied. The effects of the
planform mish divisions were investigated by performing
different computations on a swept back wing using chordwise
divisions in the range from 2 to 16 and spanwise divisions in
the range from 2 to 16. The wing geometric parameters were
studied by testing different wing configurations of aspect ratio
(AR) in the range from 6 to 12, taper ratio (TR) in the range
from 0 to 1, swept angle (SA) in the range from 0 to 60
degrees, dihedral angle (DIH) in the range from 10 to 40
degrees, angle of attack (AOA) in the range from 6 to 12
degrees, and ground effect at different values of the ratio
between the height above the ground and the chord (h/c) in the
range from 0.2 to 2.
The research was conducted to estimate the values of the
stability and control derivatives for the three-dimensional
model of multi-element lifting surface with trailing edge
control surfaces using the developed solver. The stability and
control derivatives were estimated for conventional wing
configuration aircrafts like Cessna 172, Boeing 747, and F18;
and non-conventional wing configuration aircrafts like boxwings.