ESS_Flight_Dynamics
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2-DoF-TrajectorySim
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Removed atmosisa dependence. Added automation scripts for trajectory runs.

master
Jason Ong 2 years ago
parent
8636b276c6
commit
36eceb54d7
9 changed files with 142 additions and 90 deletions
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  1. 4
      .gitignore
  2. 17
      Code/AllClean.m
  3. 14
      Code/AllRun.m
  4. 70
      Code/CalcAtmosQuantities.m
  5. 10
      Code/DragInterpolations_ICANSAT.m
  6. 8
      Code/DragInterpolations_MAD.m
  7. 8
      Code/DragInterpolations_Spark.m
  8. 18
      Code/ICANSAT/V1/ICANSAT_V1.m
  9. 83
      Code/SingleStageRocket.m

4
.gitignore
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@ -0,0 +1,4 @@
*.asv
Code/ICANSAT/V0/Result
Code/ICANSAT/V1/Result
Code/MAD/Result

17
Code/AllClean.m
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@ -0,0 +1,17 @@
mfile_name = mfilename('fullpath');
[pathstr,name,ext] = fileparts(mfile_name);
cd(pathstr);
cd ICANSAT/V0/Result
delete *.txt
fprintf("ICANSAT V0 result files deleted.\n");
cd ../../V1/Result
delete *.txt
fprintf("ICANSAT V1 result files deleted.\n");
cd ../../../MAD/Result
delete *.txt
fprintf("MAD result files deleted.\n");
cd(pathstr);

14
Code/AllRun.m
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@ -0,0 +1,14 @@
mfile_name = mfilename('fullpath');
[pathstr,name,ext] = fileparts(mfile_name);
cd(pathstr);
cd ICANSAT/V0
ICANSAT
cd ../V1
ICANSAT_V1
cd ../../MAD
MAD_Validation
cd(pathstr);

70
Code/CalcAtmosQuantities.m
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@ -0,0 +1,70 @@
function [t,p,rho] = CalcAtmosQuantities(h)
hg = h; % geometric alt in meters
R = 287; % Gas Constant
re = 6.356766e6; % radius of earth
g0 = 9.8066; % gravity
h = re/(re+hg)*hg;
% Altitude in meters
h0 = 0.0;
h1 = 11000.0;
h2 = 25000.0;
h3 = 47000;
h4 = 53000;
h5 = 79000;
h6 = 90000;
h7 = 105000;
% Temperature in Kelvin.
t0 = 288.16;
t1 = 216.66;
t2 = 216.66;
t3 = 282.66;
t4 = 282.66;
t5 = 165.66;
t6 = 165.66;
t7 = 225.66;
%%equations
a01 = (t1-t0)/(h1-h0); % slope of T/h in K/m for 0 to 11 km
a23 = (t3-t2)/(h3-h2); % slope of T/h in K/m for 25 to 47 km
a45 = (t5-t4)/(h5-h4); % slope of T/h in K/m for 53 to 79 km
a67 = (t7-t6)/(h7-h6); % slope of T/h in K/m for 90 to 105 km
% Pressure in N/m^2
p0 = 101325.0;
p1 = p0*(t1/t0)^(-g0/(a01*R));
p2 = p1*exp(-(g0/(R*t1))*(h2-h1));
p3 = p2*(t3/t2)^(-g0/(a23*R));
p4 = p3*exp(-(g0/(R*t3))*(h4-h3));
p5 = p4*(t5/t4)^(-g0/(a45*R));
p6 = p5*exp(-(g0/(R*t5))*(h6-h5));
p7 = p6*(t7/t6)^(-g0/(a67*R));
if h < h1
t = t0+a01*h;
p = p0*(t/t0)^(-g0/(a01*R));
elseif h < h2
t = t1;
p= p1*exp(-(g0/(R*t))*(h-h1));
elseif h < h3
t = t2+a23*(h-h2);
p = p2*(t/t2)^(-g0/(a23*R));
elseif h < h4
t = t3;
p = p3*exp(-(g0/(R*t))*(h-h3));
elseif h < h5
t = t4+a45*(h-h4);
p = p4*(t/t4)^(-g0/(a45*R));
elseif h < h6
t = t5;
p = p5*exp(-(g0/(R*t))*(h-h5));
elseif h < h7
t = t6+a67*(h-h6);
p = p6*(t/t6)^(-g0/(a67*R));
else
t= t7;
p = p7*exp(-(g0/(R*t))*(h-h7));
end
rho=p/(R*t); % Density in kg/m^3
end

10
Code/DragInterpolations_ICANSAT.m
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@ -1,7 +1,5 @@
function [coeffdrag] = DragInterpolations_ICANSAT(velocity, altitude, time, tb)
% global tb
drag_data = dlmread("ICANSAT/ICANSAT_DragData.txt", "\t",1,0);
% Mach No
@ -13,7 +11,13 @@ cd_full = drag_data(:,3);
% No base drag
cd_zero = drag_data(:,2);
[temp, sound_speed, P, rho] = atmosisa(altitude);
R = 8.3145;
gamma = 1.4;
M_air = 0.0289645;
[T,p,rho] = CalcAtmosQuantities(altitude);
sound_speed = sqrt(gamma*R*T/M_air);
mach_speed = velocity/sound_speed;
if (time > tb)

8
Code/DragInterpolations_MAD.m
View File

@ -27,7 +27,13 @@ cdzero_10000 = drag_data(:,3);
% Interpolate for Cd values at two stated altitudes, then interpolate for
% Cd value at a specific altitude
[temp, sound_speed, P, rho] = atmosisa(altitude);
R = 8.3145;
gamma = 1.4;
M_air = 0.0289645;
[T,p,rho] = CalcAtmosQuantities(altitude);
sound_speed = sqrt(gamma*R*T/M_air);
mach_speed = velocity/sound_speed;
% If propellant is still available

8
Code/DragInterpolations_Spark.m
View File

@ -2,7 +2,13 @@ function [coeffdrag] = DragInterpolations_Spark(velocity, altitude)
spark_drag_data = dlmread("Spark/Spark_DragData.txt", "\t",1,0);
[temp, sound_speed, P, rho] = atmosisa(altitude);
R = 8.3145;
gamma = 1.4;
M_air = 0.0289645;
[T,p,rho] = CalcAtmosQuantities(altitude);
sound_speed = sqrt(gamma*R*T/M_air);
mach_speed = velocity/sound_speed;
mach = spark_drag_data(:,1);

18
Code/ICANSAT/V1/ICANSAT_V1.m
View File

@ -5,15 +5,15 @@ addpath('../..')
% Df, Lr, Ls, ms, mp, tb
icansat = SingleStageRocket(0.1567, 2.391, 4, 14, 1.160534, 4.0);
% RunTrajectorySim(icansat, 88);
% RunTrajectorySim(icansat, 86);
% RunTrajectorySim(icansat, 84);
% RunTrajectorySim(icansat, 82);
% RunTrajectorySim(icansat, 80);
% RunTrajectorySim(icansat, 78);
% RunTrajectorySim(icansat, 76);
% RunTrajectorySim(icansat, 74);
% RunTrajectorySim(icansat, 72);
RunTrajectorySim(icansat, 88);
RunTrajectorySim(icansat, 86);
RunTrajectorySim(icansat, 84);
RunTrajectorySim(icansat, 82);
RunTrajectorySim(icansat, 80);
RunTrajectorySim(icansat, 78);
RunTrajectorySim(icansat, 76);
RunTrajectorySim(icansat, 74);
RunTrajectorySim(icansat, 72);
RunTrajectorySim(icansat, 70);
function RunTrajectorySim(rocket, launch_angle)

83
Code/SingleStageRocket.m
View File

@ -165,81 +165,9 @@ classdef SingleStageRocket < handle
obj.CalculateMassFlow(length(resampled_time));
end
function [t,p,rho] = CalcAtmosQuantities(obj, h)
hg = h; % geometric alt in meters
R = 287; % Gas Constant
re = 6.356766e6; % radius of earth
g0 = 9.8066; % gravity
h = re/(re+hg)*hg;
% Altitude in meters
h0 = 0.0;
h1 = 11000.0;
h2 = 25000.0;
h3 = 47000;
h4 = 53000;
h5 = 79000;
h6 = 90000;
h7 = 105000;
% Temperature in Kelvin.
t0 = 288.16;
t1 = 216.66;
t2 = 216.66;
t3 = 282.66;
t4 = 282.66;
t5 = 165.66;
t6 = 165.66;
t7 = 225.66;
%%equations
a01 = (t1-t0)/(h1-h0); % slope of T/h in K/m for 0 to 11 km
a23 = (t3-t2)/(h3-h2); % slope of T/h in K/m for 25 to 47 km
a45 = (t5-t4)/(h5-h4); % slope of T/h in K/m for 53 to 79 km
a67 = (t7-t6)/(h7-h6); % slope of T/h in K/m for 90 to 105 km
% Pressure in N/m^2
p0 = 101325.0;
p1 = p0*(t1/t0)^(-g0/(a01*R));
p2 = p1*exp(-(g0/(R*t1))*(h2-h1));
p3 = p2*(t3/t2)^(-g0/(a23*R));
p4 = p3*exp(-(g0/(R*t3))*(h4-h3));
p5 = p4*(t5/t4)^(-g0/(a45*R));
p6 = p5*exp(-(g0/(R*t5))*(h6-h5));
p7 = p6*(t7/t6)^(-g0/(a67*R));
if h < h1
t = t0+a01*h;
p = p0*(t/t0)^(-g0/(a01*R));
elseif h < h2
t = t1;
p= p1*exp(-(g0/(R*t))*(h-h1));
elseif h < h3
t = t2+a23*(h-h2);
p = p2*(t/t2)^(-g0/(a23*R));
elseif h < h4
t = t3;
p = p3*exp(-(g0/(R*t))*(h-h3));
elseif h < h5
t = t4+a45*(h-h4);
p = p4*(t/t4)^(-g0/(a45*R));
elseif h < h6
t = t5;
p = p5*exp(-(g0/(R*t))*(h-h5));
elseif h < h7
t = t6+a67*(h-h6);
p = p6*(t/t6)^(-g0/(a67*R));
else
t= t7;
p = p7*exp(-(g0/(R*t))*(h-h7));
end
rho=p/(R*t); % Density in kg/m^3
end
% Thrust correction based on ASTOS manual
function T = ThrustCorrection(obj, h, thrust)
[t,p,rho] = obj.CalcAtmosQuantities(h);
[t,p,rho] = CalcAtmosQuantities(h);
T = thrust*1.095 - 0.014 * p;
end
@ -277,7 +205,7 @@ classdef SingleStageRocket < handle
% Density calculation
function [T, rho] = CalculateRhoTemp(obj, h)
[T, p, rho] = obj.CalcAtmosQuantities(h);
[T,p,rho] = CalcAtmosQuantities(h);
return;
end
@ -427,15 +355,18 @@ classdef SingleStageRocket < handle
% Write data file
function WriteDataFile(obj)
if obj.mdot_constant == true
result_dat_file = obj.ID + "_Result_" + obj.launch_angle + "deg.txt";
result_dat_file = 'Result/' + obj.ID + "_Result_" + obj.launch_angle + "deg.txt";
else
text = "VariableMassFlow";
result_dat_file = obj.ID + "_Result_" + obj.launch_angle + "_deg_" + text + ".txt";
result_dat_file = 'Result/' + obj.ID + "_Result_" + obj.launch_angle + "_deg_" + text + ".txt";
end
result_matrix = [obj.t obj.x/1000 obj.y/1000 obj.M obj.thrust obj.DvDt obj.m obj.mdot obj.isp];
result_data = reshape(result_matrix, [], 9);
dlmwrite(result_dat_file, 'Time(s),x(km),y(km),Mach,Thrust(N),Acceleration(m/s2),Mass(kg),mdot(kg/s),Isp (s)','delimiter','')
dlmwrite(result_dat_file, result_data, '-append', 'delimiter', ',');
notify_text = 'Results written for ' + obj.ID + ' at ' + obj.launch_angle + " degrees.\n ";
fprintf(notify_text);
fprintf("-----------------------------------------------------------------------------------\n");
end
end
end
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