Wing Designer: LiftCoeff(gamma, panel, geo, Fu_bar, Fv_bar, Fw_bar) - File Exchange

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Highlights from
Wing Designer

InitializeGUI(airfoildata) Initializes the GUI and establishes callback functions
DeterminePanelGeometry(inputg... find coordinates for horseshoe vortices and control points and plot
DetermineProfileDrag(airfoild... Determine the wing profile drag from the airfoils' drag polars
FinalOutput(CL, CDi, CD0, y_c... calculate score and other outputs, post to the GUI
GetGeometryfromGUI(root, tip,... Get parameters from the GUI and create geometry structure "geo"
InducedDrag(gamma, geo, panel... Calculate far field induced drag
LiftCoeff(gamma, panel, geo, ... Determine the lift coefficient given the vortex strengths
PerformRegression(data) Relate the coefficients of the drag polar to Re as parabolic functions.
SpanLoading(l_s, l_t, CL, geo... Determine center of pressure and spanwise loading
StandardAtmosphere(alt) Read in altitude and output viscosity, density, and speed of sound
ZeroOutput(output) Null output to GUI when error checking is not satisfied
[C,T,A]=naca4_3d(airfoil,y,nc... determine airfoil skin and camber coordinates from NACA designation
[gamma, Fu_bar, Fv_bar, Fw_ba... Implement Vortex Lattice Method
parseNACAairfoildata parse the text file NACAdata.txt into a structure
Contents.m
Main.m
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from Wing Designer by John Rogers
Wing Designer computes aircraft performance measures from wing and engine parameters.

LiftCoeff(gamma, panel, geo, Fu_bar, Fv_bar, Fw_bar)

function [CL, u_U, v_U, w_U, l_t, l_s] = LiftCoeff(gamma, panel, geo, Fu_bar, Fv_bar, Fw_bar)
%Determine the lift coefficient given the vortex strengths
%and the influence coefficients from the vortex lattice method

u_U.c = 1/(4*pi)*Fu_bar*[gamma.c]';  %Eqn 25 in NASA paper
v_U.c = 1/(4*pi)*Fv_bar*[gamma.c]';    %Eqn 23 in NASA paper
w_U.c = 1/(4*pi)*Fw_bar*[gamma.c]';

u_U.alpha = 1/(4*pi)*Fu_bar*[gamma.alpha]';
v_U.alpha = 1/(4*pi)*Fv_bar*[gamma.alpha]';
w_U.alpha = 1/(4*pi)*Fw_bar*[gamma.alpha]';

for i = 1:geo.ns
    for j = 1:geo.nc
        if i == geo.ns  %wingtip
            DeltaGamma(i,j).c = gamma(i,j).c;
            DeltaGamma(i,j).alpha = gamma(i,j).alpha;
        elseif j == 1 %Leading edge
            DeltaGamma(i,j).c = gamma(i,j).c-gamma(i+1,j).c;
            DeltaGamma(i,j).alpha = gamma(i,j).alpha-gamma(i+1,j).alpha;
        else
            DeltaGamma(i,j).c = DeltaGamma(i,j-1).c + gamma(i,j).c - gamma(i+1,j).c;
            DeltaGamma(i,j).alpha = DeltaGamma(i,j-1).alpha + gamma(i,j).alpha - gamma(i+1,j).alpha;
        end
    end
end

l_t.c = 2/geo.S*[DeltaGamma.c]'.*[panel.cc]'.*[v_U.c];  %Eqn 24 in NASA paper
l_t.alpha = 2/geo.S*[DeltaGamma.alpha]'.*[panel.cc]'.*[v_U.alpha];  %Eqn 24 in NASA paper

l_s.c = 2/geo.S*[gamma.c]'*2.*[panel.s]'.*((ones(size(u_U))-[u_U.c])+[v_U.c].*tan([panel.sweep]'))*cos(geo.dih);  %Eqn 28 in NASA paper
l_s.alpha = 2/geo.S*[gamma.alpha]'*2.*[panel.s]'.*((ones(size(u_U))-[u_U.alpha])+[v_U.alpha].*tan([panel.sweep]'))*cos(geo.dih);  %Eqn 28 in NASA paper

% reshape([l_t.c],geo.ns,geo.nc)
% reshape([l_t.alpha],geo.ns,geo.nc)
% reshape([l_s.c],geo.ns,geo.nc)
% reshape(((ones(size(u_U))-[u_U.alpha])+[v_U.alpha].*tan([panel.sweep]')),geo.ns,geo.nc)
% reshape([panel.s],geo.ns,geo.nc)
% reshape([l_s.alpha],geo.ns,geo.nc)
% reshape([u_U.alpha],geo.ns,geo.nc)
% reshape([v_U.alpha],geo.ns,geo.nc)
% reshape([gamma.alpha],geo.ns,geo.nc)
% sum([gamma.c])
% geo.S

CL.c = 2*(sum([l_t.c])+sum([l_s.c]));
CL.alpha = 2*(sum([l_t.alpha])+sum([l_s.alpha]));

Highlights from Wing Designer

Highlights from
Wing Designer