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#openrocket_interface imports
from math import pi, log, sqrt, exp, cos, sin, radians
import os
from sys import platform as _platform
import xml.etree.ElementTree as ET # xml library
from zipfile import ZipFile
#trajectory imports
from types import SimpleNamespace
import numpy as np
import csv
import copy
#optimizer imports
from scipy.optimize import minimize, differential_evolution
import pylab
import matplotlib
import matplotlib.pyplot as plt
from mpl_toolkits.mplot3d import Axes3D
from matplotlib import rc
Gathering all input variables, design parameters, and actual constants in one place.
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global OF, rho_ipa # GLOBAL IS EVIL
# Physics
g_n = 9.80665 # kg.m/s^2 Standard gravity
R_univ = 8314.46261815324 # universal gas constant, J/ (K * kmol)
def spec_gas(M): return R_univ / M
# Chemistry, we are using LOX and IPA/H20
rho_lox = 1141.0 # kg/m^3 Density of LOX
rho_ipa = 849.28 # kg/m^3 Density of 64.8% IPA / 35.2% H20
# N2 tank parameters
prop_tank_pressure = 689476 # Pa, pressure of propellant tanks (100 PSI)
N2_temp = 293 # K, holding this constant is sketchy but easy
N2_M = 28.01 # nitrogen molecular mass [g/mol]
# (3500 psi * 2 L * 28.01 g/mol ) / (8314.46261815324 J/(K kmol) * 293 K), where P, V, T are from
# https://github.com/psas/reaction-control/blob/master/pubs/AIAA%20RCS%20Manuscript_FINAL2.pdf gives us
rcs_req_mass = 0.55492 # kg, required amount of N2 for 150 Ns stored impulse of RCS with 8 N nozzles
# Tank Materials
# https://www.aircraftspruce.com/catalog/cmpages/anh4120honeycomb01-01574.php?clickkey=5444217
# i've calculated ~395 kg/m^3 for uniform airframe density based on our density/thickness estimates below.
# measurement of LV3.1 says module is around 130 kg/m^3
rho_nomex = 48.06 # kg/m^3, density of 3/16th in overexpanded nomex honeycomb
rho_cryo = 160 # kg/m^3, density of cryogel-z insulation
rho_fiber = 1850 # kg/m^3, density of fiberglass
Al = { 'rho': 2800.0, # kg/m^3 Density
'Sy': 0.270e9} # Pa Yield strength
CF = { 'rho': 1550.0, # kg/m^3 Density
'Sy': 0.450e9} # Pa Yield strength
# Launch constants
launch_site_alt = 1401 # m, altitude of launch site above sea level
launch_tower = 60. # launch rail height in ft
min_throttle = 0.65 # the internet says between 60 - 70% is doable without ruining our lives
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# combustion gas properties ke, Re, T_ch, determined from CEArun
# with chamber pressure=350 psi, fuel temp=419.15 K,
# lox temp=90 K, OF=1.3 for fuel = 64.8% IPA (2propanol) / 35.2% H20
OF = 1.3 # O/F ratio, this is somewhat arbitrary but CEA says its good.
p_ch = 350 # chamber pressure, PSI
T_ch = 3097.82 # chamber temperature, K
ke = 1.1251 # specific heat ratio, propellant (aka gammas)
M = 23.196 # molar mass
Re = spec_gas(M) # specific gas constant, propellant
# willy wonka's fudge factors
ullage = 1.1 # percentage of length added to a tank to account for not filling
factor_of_safety = 2 # factor of safety (defunct)
mass_fudger = 1.75 # fudge factor for tank subsystems, includes contribution of tank structural lugs,
# feed system, stress concentrations, welds, slosh baffles etc.
isogrid_reduction = 1.0 # percentage less mass
airframe_fudger = 1.05 # arbitrary number to multiply all airframe+subsystem masses by
loss_factor = 1. # if < 1, then assume exhaust velocity is less than ideal (percentage)
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# airframe parameters
airframe_offset = 0.007 # m, empirical from lv3, gap between airframe inner diameter and innards
nomex_thickness = 0.00635 # m, from lv3.1 (0.25 in)
cf_thickness = 0.001905 # m, from lv3.1 (0.075 in)
airframe_thickness = nomex_thickness + cf_thickness # m, thiccness of lv3.1
min_tank_thickness = 0.001588 # m, minimum manufacturable thickness of aluminum propellant tanks (1/16th in)
insulation_layers = 2 # around LOX tank, likely need from 1-3
insulation_thickness = 0.005 # m
dist_after_f = 0.0 # m, we could place fuel tank before avionics and N2, but don't because bad idea
dist_after_o = 0.0 # m, oxygen tank still in front of feedsys, ems, engine (duh)
# upper subsystem module dimensions
nose_l = 1.50 # m
ers_l = 6 * 0.0254 # m (converted from in)
rcs_l = 6 * 0.0254 # m (converted from in)
av_l = 24 * 0.0254 # m (converted from in)
n2_l = 24 * 0.0254 # m (converted from in)
gaps = 0.050 # m, space between subsystems
# internal subsystem masses
m_nose_frame = 3.3 # nosecone outer frame weight [kg] (est from openrocket)
m_ballast = 0 # ballast weight [kg]
m_recovery = 4 # Recovery system mass estimate [kg] (lv3.1 measured at 3.5 kg)
m_rcs = 4 # guess at mass, totally made up [kg]
m_n2 = 4 # mass of n2 tank [kg] (measured)
m_avionics = 3.3 # Avionics mass [kg]
m_fins = 8.77 # total fin mass [kg], estimate from openrocket
m_ringsclamps = (.466 + 1) * 8 # weights of rings and clamps [kg]
# Engine system dimensions. 'm_' = mass, 'l_' = length
m_feed = 10.0 # kg
l_feed = 0.4572 # m, this is 18"
m_ems = 1 # kg
l_ems = 0.1016 # m, this is 4"
m_engine = 6.0 # kg
l_engine = 0.300 # m
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# SIMULATION AND OPTIMIZATION PARAMETERS
delta = 10**(-8) # a guess at a good margin for design "convergence"
mu_0 = 0.0025 # barrier parameter, this value lets altitudes approach lower bound pretty quickly
rho_0 = 2.5 # penalty parameter, i'm still playing with this value.
time_step = 0.25 # change time-step for trajectory simulation
iterations = 25 # number of escalating iteration sequences
# INITIAL DESIGN GUESS
# be sure that you start with a feasible design, otherwise the problem will be ill-conditioned
m_prop = 145. # propellant mass (kg)
mdot_0 = 3.0 # Propellant mass flow rate (kg/s)
dia = 12.65 # Rocket diameter (in)
p_e = 55. # Exit Pressure (kPa)
lower = 100. # lower and upper bounds of drag force (N) for throttling
upper = 1000.
# OPTIMIZATION CONSTRAINTS
cons_mass = 270. # GLOW constraint, kg, somewhat arbitrary
cons_ls = 22. # min launch speed from 60' tower constraint, m/s
cons_TWR = 2. # TWR constraint
cons_S_crit = 0.35 # Critical pressure ratio constraint
cons_accel = 15. # Max acceleration constraint, g's
cons_LD = 21. # L/D ratio constraint, slightly arbitrary
cons_alt = 120000. + launch_site_alt # Min altitude constraint, m (adjusted to overshoot)
cons_thrust = 8. # max ground-level thrust, kN
cons_ceiling = 150000. + launch_site_alt # base-11 maximum apogee requirement, km
can_throttle = False