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2-DoF-TrajectorySim
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Added MAD_H300 and code edits for SingleStageRocet.m

master
Jason Ong 2 years ago
parent
0d5a618d8b
commit
38d1ad1f06
10 changed files with 150 additions and 14 deletions
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  1. 1
      .gitignore
  2. 40
      Code/DragInterpolations_MAD_H300.m
  3. 2
      Code/ICANSAT/V0/ICANSAT.m
  4. 2
      Code/ICANSAT/V1/ICANSAT_V1.m
  5. 2
      Code/ICANSAT/V2/ICANSAT_V2.m
  6. 2
      Code/MAD/MAD_Validation.m
  7. 12
      Code/MAD_H300/MAD_H300.m
  8. 15
      Code/MAD_H300/MAD_H300_DragData.txt
  9. 26
      Code/MAD_H300/MAD_H300_v1_VacT_v_Time.txt
  10. 62
      Code/SingleStageRocket.m

1
.gitignore
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@ -3,3 +3,4 @@ Code/ICANSAT/V0/Result
Code/ICANSAT/V1/Result Code/ICANSAT/V1/Result
Code/ICANSAT/V2/Result Code/ICANSAT/V2/Result
Code/MAD/Result Code/MAD/Result
Code/MAD_H300/Result

40
Code/DragInterpolations_MAD_H300.m
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@ -0,0 +1,40 @@
function [coeffdrag] = DragInterpolations_MAD_H300(velocity, altitude, time, tb)
drag_data = dlmread("MAD_H300/MAD_H300_DragData.txt", "\t",1,0);
% Mach No
mach_no = drag_data(:,1);
% Altitude 0m, full base drag
cdfull_0 = drag_data(:,3);
% Altitude 30000m, full base drag
cdfull_30000 = drag_data(:,4);
% Altitude 0m, no base drag
cdzero_0 = drag_data(:,2);
% Interpolate for Cd values at two stated altitudes, then interpolate for
% Cd value at a specific 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;
% Check if propellant is still available
if (time > tb) % Propellant is not available
if (altitude > 30000)
coeffdrag = interp1(mach_no, cdfull_30000, mach_speed);
else
coeffdrag = interp1(mach_no, cdfull_0, mach_speed);
end
else % Propellant still burning
coeffdrag = interp1(mach_no, cdzero_0, mach_speed);
end
end

2
Code/ICANSAT/V0/ICANSAT.m
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@ -3,7 +3,7 @@ addpath('../')
addpath('../..') addpath('../..')
% Df, Lr, Ls, ms, mp, tb % Df, Lr, Ls, ms, mp, tb
icansat = SingleStageRocket(0.1567, 2.391, 4, 14, 1.92, 8.0);
icansat = SingleStageRocket(0.1567, 2.391, 4, 14, 1.92, 8.0, 0);
% Timestep, no. of timesteps, launch angle % Timestep, no. of timesteps, launch angle
icansat.SetSimulationParams(0.01, 10000, 88); icansat.SetSimulationParams(0.01, 10000, 88);

2
Code/ICANSAT/V1/ICANSAT_V1.m
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@ -3,7 +3,7 @@ addpath('../')
addpath('../..') addpath('../..')
% Df, Lr, Ls, ms, mp, tb % Df, Lr, Ls, ms, mp, tb
icansat = SingleStageRocket(0.1567, 2.391, 4, 14, 1.160534, 4.0);
icansat = SingleStageRocket(0.1567, 2.391, 4, 14, 1.160534, 4.0, 0);
RunTrajectorySim(icansat, 88); RunTrajectorySim(icansat, 88);
RunTrajectorySim(icansat, 86); RunTrajectorySim(icansat, 86);

2
Code/ICANSAT/V2/ICANSAT_V2.m
View File

@ -3,7 +3,7 @@ addpath('../')
addpath('../..') addpath('../..')
% Df, Lr, Ls, ms, mp, tb % Df, Lr, Ls, ms, mp, tb
icansat = SingleStageRocket(0.1567, 2.391, 4, 14, 1.160534, 3.6325);
icansat = SingleStageRocket(0.1567, 2.391, 4, 14, 1.160534, 3.6325, 0);
RunTrajectorySim(icansat, 88); RunTrajectorySim(icansat, 88);
RunTrajectorySim(icansat, 86); RunTrajectorySim(icansat, 86);

2
Code/MAD/MAD_Validation.m
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@ -1,7 +1,7 @@
clc clc
addpath('../') addpath('../')
mad = SingleStageRocket(0.205, 4.96, 10, 68, 78, 20.1);
mad = SingleStageRocket(0.205, 4.96, 10, 68, 78, 20.1, 0);
mad.SetSimulationParams(0.05, 10000, 85); mad.SetSimulationParams(0.05, 10000, 85);
mad.SetRocketID("MAD"); mad.SetRocketID("MAD");
mad.InitializeVars(); mad.InitializeVars();

12
Code/MAD_H300/MAD_H300.m
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@ -0,0 +1,12 @@
clc
addpath('../')
mad_H300 = SingleStageRocket(0.31, 5.712, 1, 148, 158, 15.77, 0.0283);
mad_H300.SetSimulationParams(0.05, 20000, 85);
mad_H300.SetRocketID("MAD_H300");
mad_H300.InitializeVars();
mad_H300.SetThrustCorrection();
mad_H300.CalculateThrust('MAD_H300_v1_VacT_v_Time.txt');
mad_H300.CalculateTrajectory();
mad_H300.WriteCompactDataFile(20);

15
Code/MAD_H300/MAD_H300_DragData.txt
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@ -0,0 +1,15 @@
Mach NBD_0K FBD_0K FBD_30K
0 0.44 0.493 0.733
0.3 0.432 0.484 0.653
0.5 0.427 0.478 0.6
0.7 0.42 0.47 0.565
0.8 0.417 0.467 0.553
0.9 0.443 0.492 0.572
0.99 0.517 0.583 0.657
1.01 0.532 0.602 0.696
1.1 0.556 0.632 0.632
1.2 0.553 0.625 0.683
1.5 0.543 0.61 0.66
2 0.458 0.513 0.56
3 0.33 0.369 0.417
4 0.244 0.269 0.331-

26
Code/MAD_H300/MAD_H300_v1_VacT_v_Time.txt
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@ -0,0 +1,26 @@
0 0
0.01 12970
0.2 29125
0.4 28002
0.6 27233
0.8 26677
1 25804
1.2 25526
1.4 25306
1.6 25123
2 24829
3 24312
3.4 24130
4 23947
5 23649
6 23385
7 23134
8 22893
9 22654
10 22411
11 22163
12 21909
13 21641
14 21366
15 21086
15.77 0

62
Code/SingleStageRocket.m
View File

@ -41,6 +41,7 @@ classdef SingleStageRocket < handle
area % Cross sectional area area % Cross sectional area
q % Dynamic pressure q % Dynamic pressure
launch_angle launch_angle
exit_area % Nozzle exit_area
% Rocket identifier (MAD, ICANSAT, Spark etc.) % Rocket identifier (MAD, ICANSAT, Spark etc.)
% For purposes of drag identification % For purposes of drag identification
@ -52,7 +53,7 @@ classdef SingleStageRocket < handle
methods methods
% Initialization % Initialization
function self = SingleStageRocket(Df, Ls, Lr, ms, mp, tb)
function self = SingleStageRocket(Df, Ls, Lr, ms, mp, tb, ae)
% Set rocket parameters % Set rocket parameters
self.Df = Df; self.Df = Df;
self.Ls = Ls; self.Ls = Ls;
@ -67,6 +68,10 @@ classdef SingleStageRocket < handle
% Derived parameters % Derived parameters
self.mt = self.ms + self.mp; self.mt = self.ms + self.mp;
% Nozzle exit area
% If ae == 0, then thrust correction using exit area is ignored.
self.exit_area = ae;
end end
% Set rocket ID % Set rocket ID
@ -166,11 +171,16 @@ classdef SingleStageRocket < handle
end end
% Thrust correction based on ASTOS manual % Thrust correction based on ASTOS manual
function T = ThrustCorrection(obj, h, thrust)
function T = AstosThrustCorrection(obj, h, thrust)
[t,p,rho] = CalcAtmosQuantities(h); [t,p,rho] = CalcAtmosQuantities(h);
T = thrust*1.095 - 0.014 * p; T = thrust*1.095 - 0.014 * p;
end end
function T = VacThrustCorrection(obj, h, thrust)
[t,p,rho] = CalcAtmosQuantities(h);
T = thrust - 0;
end
% Calculate mass at each iteration % Calculate mass at each iteration
function CalculateMassFlow(obj,n_tb) function CalculateMassFlow(obj,n_tb)
g0 = 9.8066; g0 = 9.8066;
@ -228,7 +238,10 @@ classdef SingleStageRocket < handle
case "MAD" case "MAD"
Cd = DragInterpolations_MAD(V, y, t, obj.tb); Cd = DragInterpolations_MAD(V, y, t, obj.tb);
case "MAD_H300"
Cd = DragInterpolations_MAD_H300(V, y, t, obj.tb);
% Spark data not validated % Spark data not validated
case "Spark" case "Spark"
Cd = DragInterpolations_Spark(V, y); Cd = DragInterpolations_Spark(V, y);
@ -256,9 +269,16 @@ classdef SingleStageRocket < handle
else else
% Thrust correction if boolean true % Thrust correction if boolean true
if obj.thrust_correction == true if obj.thrust_correction == true
obj.thrust(i) = obj.ThrustCorrection(obj.y(i-1), obj.thrust(i));
if obj.thrust(i) < 0
obj.thrust(i) = 0;
if obj.ID == "MAD"
obj.thrust(i) = obj.AstosThrustCorrection(obj.y(i-1), obj.thrust(i));
if obj.thrust(i) < 0
obj.thrust(i) = 0;
end
else
obj.thrust(i) = obj.VacThrustCorrection(obj.y(i-1), obj.thrust(i));
if obj.thrust(i) < 0
obj.thrust(i) = 0;
end
end end
end end
thrust_i = obj.thrust(i); thrust_i = obj.thrust(i);
@ -330,8 +350,11 @@ classdef SingleStageRocket < handle
end end
% Flight path angle in °
obj.theta = obj.theta*(180/pi); obj.theta = obj.theta*(180/pi);
obj.q = obj.q./1000;
% Dynamic pressure in kPa
obj.q = obj.q./1000;
%% PREDICTED PERFORMANCE %% PREDICTED PERFORMANCE
Vx = max(obj.V); Vx = max(obj.V);
Mx = max(obj.M); Mx = max(obj.M);
@ -360,13 +383,32 @@ classdef SingleStageRocket < handle
text = "VariableMassFlow"; text = "VariableMassFlow";
result_dat_file = 'Result/' + obj.ID + "_Result_" + obj.launch_angle + "_deg_" + text + ".txt"; result_dat_file = 'Result/' + obj.ID + "_Result_" + obj.launch_angle + "_deg_" + text + ".txt";
end 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','')
result_matrix = [obj.t obj.x/1000 obj.y/1000 obj.V obj.M obj.thrust obj.DvDt obj.m obj.mdot obj.isp obj.Cd obj.q];
result_data = reshape(result_matrix, [], 12);
dlmwrite(result_dat_file, 'Time(s),x(km),y(km),Velocity(m/s),Mach,Thrust(N),Acceleration(m/s2),Mass(kg),mdot(kg/s),Isp(s),Cd,DynPres(kPa)','delimiter','')
dlmwrite(result_dat_file, result_data, '-append', 'delimiter', ','); dlmwrite(result_dat_file, result_data, '-append', 'delimiter', ',');
notify_text = 'Results written for ' + obj.ID + ' at ' + obj.launch_angle + " degrees.\n "; notify_text = 'Results written for ' + obj.ID + ' at ' + obj.launch_angle + " degrees.\n ";
fprintf(notify_text); fprintf(notify_text);
fprintf("-----------------------------------------------------------------------------------\n"); fprintf("-----------------------------------------------------------------------------------\n");
end end
% Reduce data by n times
function WriteCompactDataFile(obj, n)
if obj.mdot_constant == true
result_dat_file = 'Result/' + obj.ID + "_Result_" + obj.launch_angle + "deg (Compact).txt";
else
text = "VariableMassFlow";
result_dat_file = 'Result/' + obj.ID + "_Result_" + obj.launch_angle + "_deg_" + text + ".txt";
end
result_matrix = [obj.t(1:n:obj.NStepEnd) obj.x(1:n:obj.NStepEnd)/1000 obj.y(1:n:obj.NStepEnd)/1000 obj.V(1:n:obj.NStepEnd) obj.M(1:n:obj.NStepEnd) ...
obj.thrust(1:n:obj.NStepEnd) obj.DvDt(1:n:obj.NStepEnd) obj.m(1:n:obj.NStepEnd) obj.mdot(1:n:obj.NStepEnd) obj.isp(1:n:obj.NStepEnd) obj.Cd(1:n:obj.NStepEnd) ...
obj.q(1:n:obj.NStepEnd)];
result_data = reshape(result_matrix, [], 12);
dlmwrite(result_dat_file, 'Time(s),x(km),y(km),Velocity(m/s),Mach,Thrust(N),Acceleration(m/s2),Mass(kg),mdot(kg/s),Isp(s),Cd,DynPres(kPa)','delimiter','')
dlmwrite(result_dat_file, result_data, '-append', 'delimiter', ',');
notify_text = 'Compact results written for ' + obj.ID + ' at ' + obj.launch_angle + " degrees.\n ";
fprintf(notify_text);
fprintf("-----------------------------------------------------------------------------------\n");
end
end end
end end
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