clear x
mPayload = 100/2.205;
sigmaUS = 0.1;
iSpECAPS = 235;
dVU = 0.3;
dBooster = 0.8;
rBooster = dBooster/2;
lNoseCone = 1.0;
rhoSolidF = 1350;
rhoECAPS = 1000;
rhoAl = 2700;
lGulfStream = 27.23;
tShell = 4e-3;
lNozzle = 0.5;
thrust0 = 30000;
pullUpAccel = 3;
alphaPullUp = 4;
sweepAngle = 45;
wingSpan = 2.0;
rootChord = 1.0;
thicknessRatio = 0.1;
tailSpan = 0.8;
tailChord = 0.5;
shellColor = [0.32 0.32 0.32];
skinThickness = 0.002;
iSpSolid = 294;
sigma = 111.3/2134.3;
[mF, mT] = SS2O( iSpECAPS, mPayload, sigmaUS, dVU );
mPayload = mPayload + mT;
h = 25000*12*.0254/1000;
m = 0.8;
a = StdAtm(h);
rEarth = 6378.165;
v = a.speedOfSound*m/1000;
mI = 2858;
hOrbit = 350;
rOrbit = hOrbit + rEarth;
vOrbit = sqrt(3.98600436e5/rOrbit);
vEarth = (2*pi/86400)*(rEarth+h);
vDrag = 1;
deltaV = vOrbit + vDrag - vEarth - v;
uE = iSpSolid*ones(1,3)*9.806/1000;
e = [sigma sigma sigma];
[pR,pRTot] = OptimalPayloadRatio( deltaV, uE, e );
[m0, mf] = EP2M( e, pR, mPayload );
mMargin = mI - m0(1);
aBooster = pi*dBooster^2/4;
mFuel = m0 - mF;
massSolid = mFuel.*(1+sigma);
lS = mFuel./(rhoSolidF*aBooster);
lU = 0.5+mF /(rhoECAPS* aBooster);
thrust = thrust0*lS(1:3)/lS(1);
mDot = thrust/(9.806*iSpSolid);
tBurn = mFuel./mDot;
lTotal = lU + lNoseCone + sum(lS);
lS = [lS+lNozzle lU];
lMax = max(lS);
yZLim = [-dBooster-wingSpan/2 dBooster+wingSpan/2];
mS = m0 - mf;
cM = sum([mS mPayload].*lS)/m0(1);
iSP = [iSpSolid iSpSolid iSpSolid iSpECAPS 0 0];
mFuel = [mFuel mF 0 0];
wingArea = WingSizingForPullUp( pullUpAccel, m, h, 2*pi, alphaPullUp*pi/180, thrust(1), m0(1), sweepAngle*pi/180 );
tailArea = 0.2*wingArea;
wingArea = 0.8*wingArea;
WingAreaToTaper( rootChord, wingArea, wingSpan, sweepAngle*pi/180 );
[taper, sweepTrailing] = WingAreaToTaper( rootChord, wingArea, wingSpan, sweepAngle*pi/180 );
[vWing, fWing] = WingFromParameters( rootChord, wingArea, wingSpan, sweepAngle*pi/180, thicknessRatio );
WingAreaToTaper( tailChord, tailArea, tailSpan, sweepAngle*pi/180 );
[vTail, fTail] = WingFromParameters( tailChord, tailArea, tailSpan, sweepAngle*pi/180, thicknessRatio );
[vRudd, fRudd] = WingFromParameters( 1.4*tailChord, 1.2*tailArea, tailSpan, sweepAngle*pi/180, thicknessRatio );
k = 1;
x{k,1} = 'Gulfstream 350 length'; x{k,2} = sprintf('%8.2f m', lGulfStream); k = k + 1;
x{k,1} = 'Gulfstream 350 payload'; x{k,2} = sprintf('%8.2f kg', mI); k = k + 1;
x{k,1} = 'Booster mass'; x{k,2} = sprintf('%8.2f kg', m0(1)); k = k + 1;
x{k,1} = 'Booster margin'; x{k,2} = sprintf('%8.2f kg', mMargin); k = k + 1;
x{k,1} = 'Booster length'; x{k,2} = sprintf('%8.2f m', lTotal); k = k + 1;
x{k,1} = 'Booster diameter'; x{k,2} = sprintf('%8.2f m', dBooster); k = k + 1;
x{k,1} = 'Delta V'; x{k,2} = sprintf('%8.2f km/s', deltaV); k = k + 1;
x{k,1} = 'Solid Isp'; x{k,2} = sprintf('%8.2f s', iSpSolid); k = k + 1;
x{k,1} = 'ECAPS Isp'; x{k,2} = sprintf('%8.2f s', iSpECAPS); k = k + 1;
x{k,1} = 'Mach separation'; x{k,2} = sprintf('%8.2f ', m); k = k + 1;
x{k,1} = 'Velocity separation'; x{k,2} = sprintf('%8.2f km/s', v); k = k + 1;
x{k,1} = 'Altitude separation'; x{k,2} = sprintf('%8.2f km', h); k = k + 1;
x{k,1} = 'Drag loss'; x{k,2} = sprintf('%8.2f km/s', vDrag); k = k + 1;
x{k,1} = 'Orbit altitude'; x{k,2} = sprintf('%8.2f km', hOrbit); k = k + 1;
x{k,1} = 'Density solid fuel'; x{k,2} = sprintf('%8.2f kg/m^3', rhoSolidF); k = k + 1;
x{k,1} = 'Density ECAPS fuel'; x{k,2} = sprintf('%8.2f kg/m^3', rhoECAPS); k = k + 1;
x{k,1} = 'ECAPS correction'; x{k,2} = sprintf('%8.2f km/s', dVU); k = k + 1;
x{k,1} = 'Burn times'; x{k,2} = sprintf('[%8.2f %8.2f %8.2f] s', tBurn); k = k + 1;
x{k,1} = 'Thrust'; x{k,2} = sprintf('[%8.1f %8.1f %8.1f] N', thrust); k = k + 1;
x{k,1} = 'Solid stage lengths'; x{k,2} = sprintf('[%8.1f %8.1f %8.1f] m', lS(1:3)); k = k + 1;
x{k,1} = 'Pull up'; x{k,2} = sprintf('%8.1f g', pullUpAccel); k = k + 1;
x{k,1} = 'Angle of attack'; x{k,2} = sprintf('%8.1f deg', alphaPullUp); k = k + 1;
x{k,1} = 'Wing sweep angle'; x{k,2} = sprintf('%8.1f deg', sweepAngle); k = k + 1;
x{k,1} = 'Wing area'; x{k,2} = sprintf('%8.2f m^2', wingArea); k = k + 1;
x{k,1} = 'Wing trailing sweep angle'; x{k,2} = sprintf('%8.2f deg', sweepTrailing*180/pi); k = k + 1;
x{k,1} = 'Wing taper'; x{k,2} = sprintf('%8.2f ', taper); k = k + 1;
x{k,1} = 'Wing root chord'; x{k,2} = sprintf('%8.2f m', rootChord); k = k + 1;
x{k,1} = 'Wing span'; x{k,2} = sprintf('%8.2f m ', wingSpan); k = k + 1;
x{k,1} = 'Center of mass'; x{k,2} = sprintf('%8.2f m ', cM); k = k + 1;
x
name = {'Stage1' 'Stage2' 'Stage3' 'Stage4' 'ConeR' 'ConeL'};
type = [2 1 1 5 3 4];
for k = 1:length(name)
BuildCADModel( 'initialize' );
BuildCADModel( 'set name' , name{k} );
BuildCADModel( 'set units', 'mks' );
m = [];
m.name = 'Core';
m.rHinge = [0;0;0];
m.bHinge.b = eye(3);
m.previousBody = [];
BuildCADModel('add body', m );
BuildCADModel( 'compute paths' );
b = [0 0 1;0 1 0;-1 0 0];
rNoseCone = (0.4/0.25)*[0.25 0.2 0.125 0];
switch type(k)
case 1
mass = massSolid(k);
m = CreateComponent( 'make', 'cylinder', 'rUpper', rBooster, 'rLower', rBooster, 'h', lS(k),'n',20, 'rA',[0;0;0],...
'name', 'Shell', 'body', 1, 'mass', mass, 'b', b, ...
'faceColor', shellColor, 'inside', 0);
BuildCADModel( 'add component', m );
[v, f] = NoseCone( rNoseCone, lNozzle, 20, 0 );
massNozzle = 2*pi*mean(rNoseCone)*lNozzle*rhoAl*skinThickness;
m = CreateComponent( 'make', 'generic', 'vertex', v, 'face', f, 'rA',[-lNozzle;0;0],...
'name', 'Nozzle', 'body', 1, 'mass', massNozzle, 'b', b, ...
'faceColor', 'black', 'inside', 0);
BuildCADModel( 'add component', m );
centerOfMass = [lS(k)/2;0;0];
inertia = Inertias( mass, [rBooster lS(k)], 'cylinder',1 );
mass = mass + massNozzle;
case 2
mass = massSolid(k);
m = CreateComponent( 'make', 'cylinder', 'rUpper', rBooster, 'rLower', rBooster, 'h', lS(k),'n',20, 'rA',[0;0;0],...
'name', 'Shell', 'body', 1, 'mass', mass, 'b', b, ...
'faceColor', shellColor, 'inside', 0);
BuildCADModel( 'add component', m );
[v, f] = NoseCone( rNoseCone, lNozzle, 20, 0 );
massNozzle = 2*pi*mean(rNoseCone)*lNozzle*rhoAl*skinThickness;
m = CreateComponent( 'make', 'generic', 'vertex', v, 'face', f, 'rA',[-lNozzle;0;0],...
'name', 'Nozzle', 'body', 1, 'mass', massNozzle, 'b', b, ...
'faceColor', 'black', 'inside', 0);
BuildCADModel( 'add component', m );
massWing = 2*wingArea*rhoAl*skinThickness;
m = CreateComponent( 'make', 'generic', 'vertex', vWing, 'face', fWing, 'rA',[cM-rootChord/2;rBooster;0],...
'name', 'Wing1', 'body', 1, 'mass', massWing, ...
'faceColor', shellColor, 'inside', 0);
BuildCADModel( 'add component', m );
m = CreateComponent( 'make', 'generic', 'vertex', vWing, 'face', fWing, 'rA',[cM-rootChord/2;-rBooster;0],...
'name', 'Wing2', 'body', 1, 'mass', massWing, 'b', [1 0 0;0 -1 0;0 0 1], ...
'faceColor',shellColor, 'inside', 0);
BuildCADModel( 'add component', m );
massTail = 2*tailArea*rhoAl*skinThickness;
m = CreateComponent( 'make', 'generic', 'vertex', vTail, 'face', fTail, 'rA',[0;rBooster;0],...
'name', 'Tail1', 'body', 1, 'mass', massTail, ...
'faceColor', shellColor, 'inside', 0);
BuildCADModel( 'add component', m );
m = CreateComponent( 'make', 'generic', 'vertex', vTail, 'face', fTail, 'rA',[0;-rBooster;0],...
'name', 'Tail2', 'body', 1, 'mass', mass, 'b', [1 0 0;0 -1 0;0 0 1], ...
'faceColor', shellColor, 'inside', 0);
BuildCADModel( 'add component', m );
m = CreateComponent( 'make', 'generic', 'vertex', vRudd, 'face', fRudd, 'rA',[0;0;rBooster],...
'name', 'Rudder', 'body', 1, 'mass', mass, 'b', [1 0 0;0 0 -1;0 1 0], ...
'faceColor', shellColor, 'inside', 0);
BuildCADModel( 'add component', m );
mass = mass + 2*massWing + massTail + massNozzle;
centerOfMass = [lS(k)/2;0;0];
inertia = Inertias( mass, [rBooster lS(k)], 'cylinder',1 );
case 5
mass = mPayload;
m = CreateComponent( 'make', 'cylinder', 'rUpper', rBooster, 'rLower', rBooster, 'h', lS(k),'n',20, 'rA',[0;0;0],...
'name', 'Shell', 'body', 1, 'mass', 1, 'b', b, ...
'faceColor', shellColor, 'inside', 0);
BuildCADModel( 'add component', m );
centerOfMass = [lS(k)/2;0;0];
inertia = Inertias( mass, [rBooster lS(k)], 'cylinder',1 );
case {3 4}
[v, f] = NoseCone( rNoseCone, lNoseCone, 20, 1 );
if( type(k) == 4 )
b = [1 0 0;0 -1 0;0 0 1]*b;
end
mass = 2*pi*mean(rNoseCone)*lNoseCone*rhoAl*skinThickness;
m = CreateComponent( 'make', 'generic', 'vertex', v, 'face', f, 'rA',[0;0;0],...
'name', 'Nose Cone', 'body', 1, 'mass', mass, 'b', b, ...
'faceColor', shellColor, 'inside', 0);
BuildCADModel( 'add component', m );
centerOfMass = [lNoseCone/2;0;0];
inertia = Inertias( mass, [rBooster lNoseCone], 'cylinder', 1 );
end
g = BuildCADModel( 'get model');
stageName = sprintf('ALASA%s',name{k});
ExportOBJ(g,stageName);
LoadCAD(sprintf('ALASA%s.obj',name{k}));
set(gca,'xlim',[-1 lMax], 'ylim', yZLim, 'zlim', yZLim );
j = 1;
z = cell(6,2);
z{j,1} = stageName; j = j + 1;
z{j,1} = 'Inertia'; z{j,2} = sprintf('[%8.2f %8.2f %8.2f; %8.2f %8.2f %8.2f; %8.2f %8.2f %8.2f]',inertia); j = j + 1;
z{j,1} = 'Mass'; z{j,2} = sprintf('%12.2f',mass);j = j + 1;
z{j,1} = 'Center of Mass'; z{j,2} = sprintf('[%8.2f %8.2f %8.2f]',centerOfMass);j = j + 1;
z{j,1} = 'Fuel Mass'; z{j,2} = sprintf('%8.2f',mFuel(k));j = j + 1;
z{j,1} = 'iSp'; z{j,2} = sprintf('%8.2f',iSP(k));
fprintf(1,'\n---------------------------------------------------------------\n');
for j = 1:6
fprintf(1,'%14s %s\n',z{j,1},z{j,2});
end
end
x =
29×2 cell array
{'Gulfstream 350 length' } {' 27.23 m' }
{'Gulfstream 350 payload' } {' 2858.00 kg' }
{'Booster mass' } {' 2005.60 kg' }
{'Booster margin' } {' 852.40 kg' }
{'Booster length' } {' 5.91 m' }
{'Booster diameter' } {' 0.80 m' }
{'Delta V' } {' 7.96 km/s' }
{'Solid Isp' } {' 294.00 s' }
{'ECAPS Isp' } {' 235.00 s' }
{'Mach separation' } {' 0.80 ' }
{'Velocity separation' } {' 0.27 km/s' }
{'Altitude separation' } {' 7.62 km' }
{'Drag loss' } {' 1.00 km/s' }
{'Orbit altitude' } {' 350.00 km' }
{'Density solid fuel' } {' 1350.00 kg/m^3' }
{'Density ECAPS fuel' } {' 1000.00 kg/m^3' }
{'ECAPS correction' } {' 0.30 km/s' }
{'Burn times' } {'[ 192.12 192.12 192.12] s'}
{'Thrust' } {'[ 30000.0 10896.9 3919.3] N'}
{'Solid stage lengths' } {'[ 3.4 1.6 0.9] m'}
{'Pull up' } {' 3.0 g' }
{'Angle of attack' } {' 4.0 deg' }
{'Wing sweep angle' } {' 45.0 deg' }
{'Wing area' } {' 1.34 m^2' }
{'Wing trailing sweep angle'} {' 18.86 deg' }
{'Wing taper' } {' 0.34 ' }
{'Wing root chord' } {' 1.00 m' }
{'Wing span' } {' 2.00 m ' }
{'Center of mass' } {' 2.51 m ' }
---------------------------------------------------------------
ALASAStage1
Inertia [ 2203.44 0.00 0.00; 0.00 2203.44 0.00; 0.00 0.00 171.20]
Mass 2139.97
Center of Mass [ 1.72 0.00 0.00]
Fuel Mass 1999.21
iSp 294.00
---------------------------------------------------------------
ALASAStage2
Inertia [ 187.53 0.00 0.00; 0.00 187.53 0.00; 0.00 0.00 61.12]
Mass 767.95
Center of Mass [ 0.79 0.00 0.00]
Fuel Mass 726.17
iSp 294.00
---------------------------------------------------------------
ALASAStage3
Inertia [ 28.92 0.00 0.00; 0.00 28.92 0.00; 0.00 0.00 21.98]
Mass 278.71
Center of Mass [ 0.44 0.00 0.00]
Fuel Mass 261.19
iSp 294.00
---------------------------------------------------------------
ALASAStage4
Inertia [ 6.05 0.00 0.00; 0.00 6.05 0.00; 0.00 0.00 7.82]
Mass 97.74
Center of Mass [ 0.26 0.00 0.00]
Fuel Mass 6.39
iSp 235.00
---------------------------------------------------------------
ALASAConeR
Inertia [ 0.96 0.00 0.00; 0.00 0.96 0.00; 0.00 0.00 0.62]
Mass 7.80
Center of Mass [ 0.50 0.00 0.00]
Fuel Mass 0.00
iSp 0.00
---------------------------------------------------------------
ALASAConeL
Inertia [ 0.96 0.00 0.00; 0.00 0.96 0.00; 0.00 0.00 0.62]
Mass 7.80
Center of Mass [ 0.50 0.00 0.00]
Fuel Mass 0.00
iSp 0.00
