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Travis Hunter 2025-04-20 12:29:36 -06:00
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commit bbe1199dbb
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.gitignore vendored
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# Prerequisites
*.d
# build directory
build/
# Compiled Object files
*.slo
*.lo

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.vscode/settings.json vendored Normal file
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{
"files.associations": {
"numeric": "cpp",
"algorithm": "cpp"
}
}

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CMakeLists.txt Normal file
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cmake_minimum_required(VERSION 3.16)
project(qtrocket LANGUAGES CXX)
# Enable testing
include(CTest)
enable_testing()
# Set C++ standard
set(CMAKE_CXX_STANDARD 17)
set(CMAKE_CXX_STANDARD_REQUIRED ON)
find_package(Python3 COMPONENTS Interpreter Development REQUIRED)
# Include directories
include_directories(include)
# Find all source files
file(GLOB_RECURSE SIM_SRC src/*.cpp)
file(GLOB_RECURSE TEST_SRC tests/*.cpp)
# Create the rocket simulator library
add_library(rocketsimlib ${SIM_SRC})
# Main example executable
add_executable(basic_flight_simulation examples/basic_flight_simulation.cpp)
target_include_directories(basic_flight_simulation
PRIVATE /usr/lib/python3.13/site-packages/numpy/_core/include/)
target_link_libraries(basic_flight_simulation PRIVATE rocketsimlib PRIVATE Python3::Python)
# Test executable
add_executable(unit_tests ${TEST_SRC})
target_link_libraries(unit_tests PRIVATE rocketsimlib)
# Register tests with CTest
add_test(NAME UnitTests COMMAND unit_tests)
# ----------------------------------------------------------------------------
# PlantUML documentation generation
# ----------------------------------------------------------------------------
find_program(PLANTUML_EXECUTABLE plantuml DOC "Path to PlantUML executable")
if(PLANTUML_EXECUTABLE)
file(GLOB UML_FILES "${CMAKE_SOURCE_DIR}/docs/*.puml")
foreach(UML_FILE ${UML_FILES})
get_filename_component(UML_NAME ${UML_FILE} NAME_WE)
set(UML_OUTPUT "${CMAKE_SOURCE_DIR}/docs/${UML_NAME}.svg")
add_custom_command(
OUTPUT ${UML_OUTPUT}
COMMAND ${PLANTUML_EXECUTABLE} -tsvg ${UML_FILE}
DEPENDS ${UML_FILE}
COMMENT "Generating UML diagram ${UML_NAME}.svg"
VERBATIM
)
list(APPEND UML_OUTPUTS ${UML_OUTPUT})
endforeach()
add_custom_target(docs_uml ALL DEPENDS ${UML_OUTPUTS})
endif()

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LICENSE
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GNU GENERAL PUBLIC LICENSE
Version 3, 29 June 2007
Copyright © 2007 Free Software Foundation, Inc. <https://fsf.org/>
Everyone is permitted to copy and distribute verbatim copies of this license document, but changing it is not allowed.
Preamble
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END OF TERMS AND CONDITIONS
How to Apply These Terms to Your New Programs
If you develop a new program, and you want it to be of the greatest possible use to the public, the best way to achieve this is to make it free software which everyone can redistribute and change under these terms.
To do so, attach the following notices to the program. It is safest to attach them to the start of each source file to most effectively state the exclusion of warranty; and each file should have at least the “copyright” line and a pointer to where the full notice is found.
qtrocket2
Copyright (C) 2025 travis
This program is free software: you can redistribute it and/or modify it under the terms of the GNU General Public License as published by the Free Software Foundation, either version 3 of the License, or (at your option) any later version.
This program is distributed in the hope that it will be useful, but WITHOUT ANY WARRANTY; without even the implied warranty of MERCHANTABILITY or FITNESS FOR A PARTICULAR PURPOSE. See the GNU General Public License for more details.
You should have received a copy of the GNU General Public License along with this program. If not, see <https://www.gnu.org/licenses/>.
Also add information on how to contact you by electronic and paper mail.
If the program does terminal interaction, make it output a short notice like this when it starts in an interactive mode:
qtrocket2 Copyright (C) 2025 travis
This program comes with ABSOLUTELY NO WARRANTY; for details type `show w'.
This is free software, and you are welcome to redistribute it under certain conditions; type `show c' for details.
The hypothetical commands `show w' and `show c' should show the appropriate parts of the General Public License. Of course, your program's commands might be different; for a GUI interface, you would use an “about box”.
You should also get your employer (if you work as a programmer) or school, if any, to sign a “copyright disclaimer” for the program, if necessary. For more information on this, and how to apply and follow the GNU GPL, see <https://www.gnu.org/licenses/>.
The GNU General Public License does not permit incorporating your program into proprietary programs. If your program is a subroutine library, you may consider it more useful to permit linking proprietary applications with the library. If this is what you want to do, use the GNU Lesser General Public License instead of this License. But first, please read <https://www.gnu.org/philosophy/why-not-lgpl.html>.

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# ? Problem Statement: Rocket Design and Flight Simulation Application
## Overview
There is a critical need for an intuitive, extensible, and technically rigorous software application
that empowers hobbyists, students, certification candidates, and early-career engineers to design,
simulate, and optimize low-power, mid-power, high-power, and amateur-class rockets. The tool must
emphasize accessibility, modularity, and educational transparency, and be written in clean,
modern C++ using an object-oriented architecture. It must support both immediate practical needs
(rocket design, flight prediction) and deeper study of aerospace physics and software engineering
principles.
The platform must remain open-source, support .ork file compatibility, and operate reliably across
Linux, Mac, and Windows environments.
## Goals
- **Modern Design Tools**: Allow users to easily create and modify rocket models, specifying
parameters like airframes, mass properties, center of gravity (CG), center of pressure (CP),
propulsion systems, and aerodynamic surfaces.
- **Physics-Based Flight Simulation**: Simulate rocket flight through launch, coast, apogee, descent,
and recovery, using accurate force models (thrust, drag, gravity, weather effects).
- **Extensible Simulation Architecture**: Architect the system to initially support
3 Degree-of-Freedom (3-DoF) simulations, but natively prepare for future extension to 6-DoF
(full 3D translation and rotation dynamics) without major rework.
- **Educational Transparency**: Build the codebase to be highly readable, logically organized, and
deeply documented, promoting learning about flight dynamics and systems modeling.
- **Component Modularity**: Each rocket component (e.g., motors, fins, payloads, recovery systems)
must exist as independent, interchangeable modules.
- **Visualization**: Provide meaningful visual outputs, including:
- 2D/3D trajectory plots
- Stability margin graphs (e.g., CG-CP margin over time)
- Thrust, velocity, and altitude vs. time graphs
- **Certification-Ready Fidelity**: Achieve simulation fidelity that can support Tripoli Rocketry
Association and NAR Level 3 certification requirements.
- **Competitive Capability**: Aim to function as a drop-in replacement or superior alternative to
tools like OpenRocket and RockSim Pro.
- **Cross-Platform Support**: Deliver fully supported builds on Linux, Mac, and Windows.
## User Personas
- **Hobbyist Rocketeer**: Designs and refines personal rockets, needing accuracy and usability
without extensive technical setup.
- **High School/University Student**: Builds rockets for courses, competitions, or research projects,
using simulation to test and verify designs.
- **STEM Educator**: Leverages the tool in classrooms to teach core concepts of dynamics, propulsion,
aerodynamics, and systems engineering.
- **Certification Candidate (NAR/Tripoli)**: Designs rockets intended for Level 1, 2, or 3
certification and needs trustworthy simulation results.
- **Aerospace Engineering Student/Professional**: Uses the tool to prototype amateur designs, and
values modular, clear, and modern C++ code.
## Operating Constraints
- **Language**: Must be written in modern C++ (C++17 or newer) with strong attention to modularity,
memory safety, and performance.
- **Portability**: Must fully support Linux, Mac, and Windows platforms with minimal
platform-specific code.
- **Open Source**: Must be licensed under a permissive open-source license (e.g., MIT or BSD
3-clause) to encourage adoption, study, and contribution.
- **File Format Compatibility**: Must support import/export of OpenRocket (.ork) files, ensuring
interoperability with existing hobbyist ecosystems.
- **Performance**: Must run smoothly on consumer-grade hardware and support both interactive design
work and batch simulation runs for optimization.
- **Incremental Upgrades**: Initial release will support 3-DoF dynamics (translational motion only)
but architecture must cleanly allow extension to full 6-DoF dynamics (translational + rotational
motion).
## User Interface Requirements
- **Modular Design**: The UI should be modular, allowing for easy addition of new features and
components without breaking existing functionality.
- **Intuitive Interaction**: The interface should be intuitive and user-friendly, with clear
labels, tooltips, and responsive design.

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@startuml
participant "FlightSimulator" as FS
participant "Rocket" as Rocket
participant "Stage" as Stage
participant "Motor" as Motor
participant "ForcesModel" as FM
participant "Environment" as Env
participant "Integrator" as Integrator
participant "FlightState" as State
participant "RecoverySystem" as Recovery
== Launch Preparation ==
FS -> Rocket: prepareForFlight(environment)
== Launch Simulation ==
loop simulation time steps
FS -> Stage: updateMotors(deltaTime)
FS -> FM: computeNetForce(State)
FM -> Rocket: query mass/CG
FM -> Env: get air density, gravity
FM -> Motor: get thrust
FS -> Integrator: step(State, netForce, deltaTime)
Integrator -> Rocket: query mass
FS -> Rocket: applyFlightState(State)
alt FS -> Rocket: check for burnout, apogee, recovery deployment
Rocket -> RecoverySystem: checkDeploymentCondition
end
== Landing ==
FS -> FlightState: detect low altitude and low vertical velocity
note over FS
End simulation after landing detected
end note
@enduml

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@startuml
abstract class Plotter {
{abstract} +plot2D()
{abstract} +addMarker()
{abstract} +plotPosVelAcc()
}
class MatplotlibPlotter {
+plot2D()
+addMarker()
+plotPosVelAcc()
}
Plotter <|-- MatplotlibPlotter
@enduml

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@startuml
class Rocket {
+addStage()
+updateMassProperties()
+prepareForFlight()
+applyFlightState()
+getTotalMass()
+getCenterOfGravity()
+getCenterOfPressure()
}
class Stage {
+addMotor()
+setAirframe()
+setFinSet()
+setRecoverySystem()
+updateMassProperties()
+getTotalThrust()
}
class Motor {
+addThrustDataPoint()
+ignite()
+update()
+getCurrentThrust()
}
class Airframe {
+getLength()
+getReferenceArea()
}
class FinSet {
+calculateNormalForceCoefficient()
+calculateCenterOfPressure()
+calculateDragArea()
+calculateMass()
}
class RecoverySystem {
+checkDeploymentCondition()
+deploy()
+isDeployed()
}
Rocket "1" *-- "*" Stage
Stage "1" *-- "*" Motor
Stage "1" *-- "0..1" Airframe
Stage "1" *-- "0..1" FinSet
Stage "1" *-- "0..1" RecoverySystem
@enduml

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@startuml
class FlightSimulator {
+run()
+step()
+initialize()
+handleEvents()
+checkTermination()
+updateMotors()
}
class ForcesModel {
+computeNetForce()
+computeNetMoment()
+computeThrust()
+computeAerodynamicDrag()
+computeGravity()
}
class FlightState {
+getPosition()
+setPosition()
+getVelocity()
+setVelocity()
+getAcceleration()
+setAcceleration()
+getOrientation()
+setOrientation()
+getAngularVelocity()
+setAngularVelocity()
+getTime()
+setTime()
}
class Environment {
+getAirDensity()
+getGravity()
}
class Integrator {
+step()
+eulerIntegration()
}
FlightSimulator --> ForcesModel
FlightSimulator --> Integrator
FlightSimulator --> FlightState
FlightSimulator --> Environment
ForcesModel --> Rocket
ForcesModel --> Environment
Integrator --> Rocket
Integrator --> ForcesModel
@enduml

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// examples/basic_flight_simulation.cpp
#include <iostream>
#include <memory>
// Include headers
#include "Rocket.h"
#include "Stage.h"
#include "Motor.h"
#include "Airframe.h"
#include "FinSet.h"
#include "RecoverySystem.h"
#include "FlightSimulator.h"
#include "Environment.h"
#include "FlightState.h"
#include "ForcesModel.h"
#include "Integrator.h"
#include "plotting/MatplotlibPlotter.h"
int main() {
// Create environment: sea level, 15?C
auto environment = std::make_shared<Environment>(
101325.0, // surface pressure (Pa)
288.15, // surface temperature (K)
0.0 // launch altitude (m)
);
// Build Rocket
auto rocket = std::make_shared<Rocket>("Test Rocket");
// Build Stage
auto stage = std::make_unique<Stage>("First Stage");
// Add Motor
auto motor = std::make_unique<Motor>("F50",
0.050, // 50 grams of propellant
50.0 // total impulse (Ns)
);
// Simple thrust curve (linear ramp-up and down)
motor->addThrustDataPoint(0.0, 0.0);
motor->addThrustDataPoint(0.1, 50.0);
motor->addThrustDataPoint(1.5, 50.0);
motor->addThrustDataPoint(1.6, 0.0);
stage->addMotor(std::move(motor));
// Add Airframe
auto airframe = std::make_unique<Airframe>("Standard Tube",
1.0, // length (m)
0.1, // diameter (m)
1.2, // dry mass (kg)
0.75 // drag coefficient (unitless)
);
stage->setAirframe(std::move(airframe));
// Add FinSet
auto finset = std::make_unique<FinSet>("Tri-Fins",
3, // number of fins
0.15, // root chord (m)
0.10, // tip chord (m)
0.10, // span (m)
0.05, // sweep length (m)
0.003, // thickness (m)
500.0 // material density (kg/m3)
);
stage->setFinSet(std::move(finset));
// Add Recovery System
auto recovery = std::make_unique<RecoverySystem>("Main Parachute",
RecoverySystem::DeploymentType::Apogee,
0.0, // trigger value not needed for apogee deploy
1.5, // drag coefficient (very large for parachutes)
1.0 // reference area (m2)
);
stage->setRecoverySystem(std::move(recovery));
// Finalize Rocket
rocket->addStage(std::move(stage));
rocket->updateMassProperties();
// Setup Simulation
FlightSimulator simulator(rocket, environment);
// Run simulation
simulator.run(30.0, 0.01); // 30 seconds max, 10ms timestep
// Output simple telemetry
const auto& flightLog = simulator.getFlightLog();
MatplotlibPlotter plotter;
// Build flight data vectors
std::vector<double> time, altitude, velocity, acceleration;
for (const auto& state : flightLog) {
time.push_back(state.getTime());
altitude.push_back(state.getPosition()[2]);
velocity.push_back(state.getVelocity()[2]);
acceleration.push_back(state.getAcceleration()[2]);
}
// --- Detect Apogee ---
double maxAltitude = 0.0;
double apogeeTime = 0.0;
for (const auto& state : flightLog) {
double alt = state.getPosition()[2];
if (alt > maxAltitude) {
maxAltitude = alt;
apogeeTime = state.getTime();
}
}
plotter.addMarker(apogeeTime, "Apogee");
// --- Detect Motor Burnout ---
// (Simple version: find when thrust becomes near zero)
bool burnoutDetected = false;
for (size_t i = 0; i < flightLog.size(); ++i) {
if (i == 0) continue; // skip first entry
double velPrev = flightLog[i-1].getVelocity()[2];
double velCurr = flightLog[i].getVelocity()[2];
// If velocity was increasing and then starts decreasing significantly
if (velPrev > velCurr + 1.0) { // "significant" drop
plotter.addMarker(flightLog[i].getTime(), "Motor Burnout (inferred)");
burnoutDetected = true;
break;
}
}
if (!burnoutDetected) {
std::cout << "Warning: Motor burnout not detected automatically.\n";
}
// --- Plot Altitude vs Time ---
plotter.plot2D(time, altitude, "Altitude vs Time", "Time (s)", "Altitude (m)");
// --- Plot Velocity vs Time ---
plotter.plot2D(time, velocity, "Velocity vs Time", "Time (s)", "Velocity (m/s)");
// --- Plot Acceleration vs Time ---
plotter.plot2D(time, acceleration, "Acceleration vs Time", "Time (s)", "Acceleration (m/s2)");
plotter.plotPosVelAcc(time, altitude, velocity, acceleration);
return 0;
}

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#ifndef AIRFRAME_H
#define AIRFRAME_H
#include <string>
/**
* @brief Represents the structural body and aerodynamic shell of a rocket stage.
*
* The Airframe class manages geometric dimensions, mass properties,
* and aerodynamic characteristics needed for simulation.
*/
class Airframe {
public:
/**
* @brief Constructs a new Airframe.
* @param name Name of the airframe component.
* @param length Total length of the airframe (meters).
* @param diameter Maximum body diameter (meters).
* @param dryMass Dry mass of the airframe structure (kilograms).
* @param dragCoefficient Estimated drag coefficient (unitless).
*/
Airframe(const std::string& name,
double length,
double diameter,
double dryMass,
double dragCoefficient);
/**
* @brief Default destructor.
*/
~Airframe() = default;
/**
* @brief Gets the total length of the airframe.
* @return Length in meters.
*/
double getLength() const;
/**
* @brief Gets the maximum diameter of the airframe.
* @return Diameter in meters.
*/
double getDiameter() const;
/**
* @brief Gets the dry mass of the airframe.
* @return Mass in kilograms.
*/
double getDryMass() const;
/**
* @brief Gets the aerodynamic drag coefficient.
* @return Drag coefficient (dimensionless).
*/
double getDragCoefficient() const;
/**
* @brief Calculates the reference area for drag computation.
*
* Reference area = Pi * (diameter / 2)^2
*
* @return Cross-sectional area in square meters.
*/
double getReferenceArea() const;
/**
* @brief Gets the name of the airframe.
* @return Name as a constant reference.
*/
const std::string& getName() const;
private:
std::string name_; ///< Name of the airframe.
double length_; ///< Total length [m].
double diameter_; ///< Maximum diameter [m].
double dryMass_; ///< Structural dry mass [kg].
double dragCoefficient_; ///< Drag coefficient (Cd).
};
#endif // AIRFRAME_H

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#ifndef ENVIRONMENT_H
#define ENVIRONMENT_H
/**
* @brief Models atmospheric and gravitational conditions for the flight simulation.
*
* The Environment class provides atmospheric properties (air density, pressure, temperature)
* and gravity based on altitude. Future versions may support wind models.
*/
class Environment {
public:
/**
* @brief Constructs a new Environment.
* @param surfacePressure Surface pressure at launch site (Pascals).
* @param surfaceTemperature Surface temperature at launch site (Kelvin).
* @param launchAltitude Launch site altitude (meters above sea level).
*/
Environment(double surfacePressure,
double surfaceTemperature,
double launchAltitude);
/**
* @brief Default destructor.
*/
~Environment() = default;
/**
* @brief Returns the atmospheric density at a given altitude.
* @param altitude Altitude above sea level (meters).
* @return Air density in kg/m3.
*/
double getAirDensity(double altitude) const;
/**
* @brief Returns the gravitational acceleration at a given altitude.
* @param altitude Altitude above sea level (meters).
* @return Gravitational acceleration in m/s2.
*/
double getGravity(double altitude) const;
/**
* @brief Returns the surface pressure at launch site.
* @return Pressure in Pascals.
*/
double getSurfacePressure() const;
/**
* @brief Returns the surface temperature at launch site.
* @return Temperature in Kelvin.
*/
double getSurfaceTemperature() const;
/**
* @brief Returns the launch altitude above sea level.
* @return Altitude in meters.
*/
double getLaunchAltitude() const;
private:
double surfacePressure_; ///< Surface pressure at launch site [Pa].
double surfaceTemperature_; ///< Surface temperature at launch site [K].
double launchAltitude_; ///< Launch altitude above sea level [m].
/**
* @brief Internal method to compute standard atmosphere temperature at altitude.
* @param altitude Altitude above sea level (meters).
* @return Temperature at altitude (Kelvin).
*/
double getStandardTemperature(double altitude) const;
/**
* @brief Internal method to compute standard atmosphere pressure at altitude.
* @param altitude Altitude above sea level (meters).
* @return Pressure at altitude (Pascals).
*/
double getStandardPressure(double altitude) const;
};
#endif // ENVIRONMENT_H

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#ifndef FINSET_H
#define FINSET_H
#include <string>
/**
* @brief Represents a set of aerodynamic fins attached to a rocket stage.
*
* The FinSet class models stability contributions, drag effects,
* and center of pressure location based on fin geometry and number.
*/
class FinSet {
public:
/**
* @brief Constructs a new FinSet.
* @param name Name of the fin set.
* @param numberOfFins Number of fins in the set.
* @param rootChord Root chord length (meters).
* @param tipChord Tip chord length (meters).
* @param span Fin span (meters).
* @param sweepLength Sweep length of the fin (meters).
* @param thickness Fin thickness (meters).
* @param materialDensity Density of fin material (kg/m3).
*/
FinSet(const std::string& name,
int numberOfFins,
double rootChord,
double tipChord,
double span,
double sweepLength,
double thickness,
double materialDensity);
/**
* @brief Default destructor.
*/
~FinSet() = default;
/**
* @brief Calculates the normal force coefficient contribution of the fin set.
* @return Normal force coefficient (Cn).
*/
double calculateNormalForceCoefficient() const;
/**
* @brief Calculates the location of the center of pressure contribution of the fin set.
* @return Distance from leading edge of root chord (meters).
*/
double calculateCenterOfPressure() const;
/**
* @brief Calculates the drag area contribution of the fins.
* @return Drag area (Cd x A) in square meters.
*/
double calculateDragArea() const;
/**
* @brief Calculates the total mass of the fin set.
* @return Mass in kilograms.
*/
double calculateMass() const;
/**
* @brief Gets the name of the fin set.
* @return Fin set name as a constant reference.
*/
const std::string& getName() const;
private:
std::string name_; ///< Name of the fin set.
int numberOfFins_; ///< Number of fins in the set.
double rootChord_; ///< Root chord length [m].
double tipChord_; ///< Tip chord length [m].
double span_; ///< Span (distance from airframe to tip) [m].
double sweepLength_; ///< Sweep length (leading edge sweep) [m].
double thickness_; ///< Thickness of fin [m].
double materialDensity_; ///< Material density [kg/m3].
/**
* @brief Helper method to calculate the planform area of a single fin.
* @return Planform area in square meters.
*/
double calculateSingleFinArea() const;
};
#endif // FINSET_H

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#ifndef FLIGHTSIMULATOR_H
#define FLIGHTSIMULATOR_H
#include <memory>
#include <vector>
class Rocket;
class Environment;
class ForcesModel;
class FlightState;
class Integrator;
/**
* @brief Manages the simulation of a rocket flight from launch to landing.
*
* The FlightSimulator coordinates the rocket, environment, force models,
* and numerical integration to simulate rocket flight dynamics over time.
*/
class FlightSimulator {
public:
/**
* @brief Constructs a new FlightSimulator.
* @param rocket Pointer to the rocket to simulate.
* @param environment Pointer to the launch environment.
*/
FlightSimulator(std::shared_ptr<Rocket> rocket,
std::shared_ptr<Environment> environment);
/**
* @brief Default destructor.
*/
~FlightSimulator() = default;
/**
* @brief Runs the full flight simulation.
* @param maxSimulationTime Maximum allowable simulation time (seconds).
* @param timeStep Initial time step for integration (seconds).
*/
void run(double maxSimulationTime, double timeStep);
/**
* @brief Returns the recorded flight states over time.
* @return Vector of FlightState snapshots.
*/
const std::vector<FlightState>& getFlightLog() const;
private:
std::shared_ptr<Rocket> rocket_; ///< Rocket being simulated.
std::shared_ptr<Environment> environment_; ///< Atmospheric and gravity conditions.
std::shared_ptr<ForcesModel> forcesModel_; ///< Computes forces and moments on the rocket.
std::unique_ptr<Integrator> integrator_; ///< Integrates equations of motion.
std::vector<FlightState> flightLog_; ///< Time history of flight state snapshots.
bool hasLaunched_ = false; ///< Have we left the lauch pad/rail?
/**
* @brief Initializes simulation (prepare rocket, set initial conditions).
*/
void initialize();
/**
* @brief Advances the simulation by one time step.
* @param deltaTime The time step size (seconds).
*/
void step(double deltaTime);
/**
* @brief Detects and processes key events (burnout, separation, recovery).
* @param state Current flight state.
*/
void handleEvents(FlightState& state);
/**
* @brief Checks if simulation termination conditions are met (e.g., landed).
* @param state Current flight state.
* @return True if simulation should stop.
*/
bool checkTermination(const FlightState& state);
void updateMotors(double deltaTime);
};
#endif // FLIGHTSIMULATOR_H

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#ifndef FLIGHTSTATE_H
#define FLIGHTSTATE_H
#include <array>
/**
* @brief Represents the physical state of the rocket at a given simulation time.
*
* The FlightState contains position, velocity, acceleration, orientation,
* and angular velocity information, future-proofed for 6-DoF dynamics.
*/
class FlightState {
public:
/**
* @brief Default constructor. Initializes to zero state.
*/
FlightState();
/**
* @brief Gets the current position vector.
* @return Position (x, y, z) in meters.
*/
const std::array<double, 3>& getPosition() const;
/**
* @brief Sets the current position vector.
* @param pos Position (x, y, z) in meters.
*/
void setPosition(const std::array<double, 3>& pos);
/**
* @brief Gets the current velocity vector.
* @return Velocity (vx, vy, vz) in meters per second.
*/
const std::array<double, 3>& getVelocity() const;
/**
* @brief Sets the current velocity vector.
* @param vel Velocity (vx, vy, vz) in meters per second.
*/
void setVelocity(const std::array<double, 3>& vel);
/**
* @brief Gets the current acceleration vector.
* @return Acceleration (ax, ay, az) in meters per second squared.
*/
const std::array<double, 3>& getAcceleration() const;
/**
* @brief Sets the current acceleration vector.
* @param acc Acceleration (ax, ay, az) in meters per second squared.
*/
void setAcceleration(const std::array<double, 3>& acc);
/**
* @brief Gets the current orientation quaternion.
* @return Orientation quaternion (w, x, y, z).
*/
const std::array<double, 4>& getOrientation() const;
/**
* @brief Sets the current orientation quaternion.
* @param quat Orientation quaternion (w, x, y, z).
*/
void setOrientation(const std::array<double, 4>& quat);
/**
* @brief Gets the current angular velocity vector.
* @return Angular velocity (roll rate, pitch rate, yaw rate) in radians per second.
*/
const std::array<double, 3>& getAngularVelocity() const;
/**
* @brief Sets the current angular velocity vector.
* @param angVel Angular velocity (roll rate, pitch rate, yaw rate) in radians per second.
*/
void setAngularVelocity(const std::array<double, 3>& angVel);
/**
* @brief Gets the elapsed simulation time.
* @return Elapsed time in seconds.
*/
double getTime() const;
/**
* @brief Sets the elapsed simulation time.
* @param time Time in seconds.
*/
void setTime(double time);
private:
std::array<double, 3> position_; ///< (x, y, z) position in meters.
std::array<double, 3> velocity_; ///< (vx, vy, vz) velocity in m/s.
std::array<double, 3> acceleration_; ///< (ax, ay, az) acceleration in m/s2.
std::array<double, 4> orientation_; ///< Orientation quaternion (w, x, y, z).
std::array<double, 3> angularVelocity_; ///< Angular rates (roll, pitch, yaw) in rad/s.
double time_; ///< Elapsed simulation time [s].
};
#endif // FLIGHTSTATE_H

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#ifndef FORCESMODEL_H
#define FORCESMODEL_H
#include <array>
#include <memory>
class Rocket;
class Environment;
class FlightState;
/**
* @brief Computes aerodynamic, thrust, and gravitational forces acting on the rocket.
*
* The ForcesModel provides force and moment outputs based on the current rocket state,
* environmental conditions, and rocket configuration.
*/
class ForcesModel {
public:
/**
* @brief Constructs a new ForcesModel.
* @param rocket Pointer to the rocket configuration.
* @param environment Pointer to the simulation environment.
*/
ForcesModel(std::shared_ptr<Rocket> rocket,
std::shared_ptr<Environment> environment);
/**
* @brief Default destructor.
*/
~ForcesModel() = default;
/**
* @brief Computes the net external force vector acting on the rocket.
* @param state Current flight state.
* @return Force vector (Fx, Fy, Fz) in Newtons.
*/
std::array<double, 3> computeNetForce(const FlightState& state) const;
/**
* @brief Computes the net external moment vector acting on the rocket.
*
* Initially returns zero (for 3-DoF), but structured for 6-DoF extension.
*
* @param state Current flight state.
* @return Moment vector (Mx, My, Mz) in Newton-meters.
*/
std::array<double, 3> computeNetMoment(const FlightState& state) const;
private:
std::shared_ptr<Rocket> rocket_; ///< Rocket model reference.
std::shared_ptr<Environment> environment_; ///< Atmospheric and gravity conditions.
/**
* @brief Computes aerodynamic drag force based on velocity and rocket configuration.
* @param state Current flight state.
* @return Drag force vector (Newton).
*/
std::array<double, 3> computeAerodynamicDrag(const FlightState& state) const;
/**
* @brief Computes thrust force based on motor outputs.
* @param state Current flight state.
* @return Thrust force vector (Newton).
*/
std::array<double, 3> computeThrust(const FlightState& state) const;
/**
* @brief Computes gravitational force based on altitude and mass.
* @param state Current flight state.
* @return Gravity force vector (Newton).
*/
std::array<double, 3> computeGravity(const FlightState& state) const;
};
#endif // FORCESMODEL_H

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#ifndef INTEGRATOR_H
#define INTEGRATOR_H
#include <memory>
#include <array>
class Rocket;
class Environment;
class ForcesModel;
class FlightState;
/**
* @brief Advances the rocket's flight state using numerical integration.
*
* The Integrator uses simple time-stepping methods (e.g., Euler, RK4) to
* update position, velocity, and optionally orientation and angular velocity.
*/
class Integrator {
public:
/**
* @brief Constructs a new Integrator.
* @param rocket Pointer to the rocket being simulated.
* @param forcesModel Pointer to the forces model.
*/
Integrator(std::shared_ptr<Rocket> rocket,
std::shared_ptr<ForcesModel> forcesModel);
/**
* @brief Default destructor.
*/
~Integrator() = default;
/**
* @brief Advances the flight state forward by one time step.
* @param state The flight state to update.
* @param deltaTime Time step size (seconds).
*/
void step(FlightState& state, const std::array<double, 3>& netForce, double deltaTime);
private:
std::shared_ptr<Rocket> rocket_; ///< Rocket model reference.
std::shared_ptr<ForcesModel> forcesModel_; ///< Forces model reference.
/**
* @brief Performs simple Euler integration for translational motion.
* @param state Current flight state.
* @param netForce Force vector acting on rocket [N].
* @param deltaTime Time step size (seconds).
*/
void eulerIntegration(FlightState& state,
const std::array<double, 3>& netForce,
double deltaTime);
};
#endif // INTEGRATOR_H

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#ifndef MOTOR_H
#define MOTOR_H
#include <string>
#include <vector>
#include <utility> // for std::pair
/**
* @brief Represents a rocket motor with thrust characteristics over time.
*
* The Motor class manages the thrust curve, ignition timing, burn duration,
* and remaining propellant mass of a single propulsion unit.
*/
class Motor {
public:
/**
* @brief Constructs a new Motor with a given name and mass properties.
* @param name Name of the motor.
* @param initialPropellantMass Initial mass of propellant (kg).
* @param totalImpulse Total impulse (Ns) for basic verification.
*/
Motor(const std::string& name, double initialPropellantMass, double totalImpulse);
/**
* @brief Default destructor.
*/
~Motor() = default;
/**
* @brief Adds a thrust curve data point.
*
* Should be called during setup. Time must be monotonically increasing.
*
* @param time Time since ignition (seconds).
* @param thrust Thrust at this time (Newtons).
*/
void addThrustDataPoint(double time, double thrust);
/**
* @brief Starts motor ignition (sets internal ignition time to zero).
*/
void ignite();
/**
* @brief Updates the motor status based on elapsed time.
* @param deltaTime Time step in seconds.
*/
void update(double deltaTime);
/**
* @brief Gets the current thrust output.
* @return Current thrust in Newtons.
*/
double getCurrentThrust() const;
/**
* @brief Checks whether the motor has burned out.
* @return True if burnout has occurred.
*/
bool isBurnedOut() const;
/**
* @brief Returns the remaining propellant mass.
* @return Remaining propellant mass in kilograms.
*/
double getRemainingPropellantMass() const;
/**
* @brief Gets the name of the motor.
* @return Motor name as a constant reference.
*/
const std::string& getName() const;
private:
std::string name_; ///< Name of the motor.
double initialPropellantMass_; ///< Initial propellant mass [kg].
double remainingPropellantMass_; ///< Current propellant mass [kg].
double totalImpulse_; ///< Total impulse [Ns], used for sanity checks.
std::vector<std::pair<double, double>> thrustCurve_; ///< Thrust curve (time, thrust) points.
double ignitionTime_; ///< Elapsed time since ignition [s].
bool ignited_; ///< Whether the motor has been ignited.
bool burnedOut_; ///< Whether the motor has finished burning.
double currentThrust_; ///< Cached current thrust [N].
/**
* @brief Linearly interpolates thrust from the thrust curve.
* @param time Time since ignition (seconds).
* @return Interpolated thrust (Newtons).
*/
double interpolateThrust(double time) const;
};
#endif // MOTOR_H

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#ifndef RECOVERYSYSTEM_H
#define RECOVERYSYSTEM_H
#include <string>
/**
* @brief Represents a recovery system for a rocket stage.
*
* The RecoverySystem class models parachutes, drogues, or streamers
* and handles deployment logic based on flight conditions.
*/
class RecoverySystem {
public:
/**
* @brief Deployment trigger types.
*/
enum class DeploymentType {
Apogee,
Altitude,
Timer
};
/**
* @brief Constructs a new RecoverySystem.
* @param name Name of the recovery device.
* @param deploymentType Type of deployment trigger (apogee, altitude, timer).
* @param triggerValue Value associated with deployment trigger (e.g., altitude in meters or time in seconds).
* @param dragCoefficient Drag coefficient of the recovery device.
* @param referenceArea Reference area (canopy projected area) in square meters.
*/
RecoverySystem(const std::string& name,
DeploymentType deploymentType,
double triggerValue,
double dragCoefficient,
double referenceArea);
/**
* @brief Default destructor.
*/
~RecoverySystem() = default;
/**
* @brief Checks if deployment conditions are met based on flight state.
* @param altitude Current altitude above ground (meters).
* @param velocity Current vertical velocity (m/s).
* @param time Elapsed flight time (seconds).
* @param atApogee Flag indicating if rocket has reached apogee.
* @return True if deployment should occur.
*/
bool checkDeploymentCondition(double altitude, double velocity, double time, bool atApogee) const;
/**
* @brief Marks the recovery system as deployed.
*/
void deploy();
/**
* @brief Returns whether the recovery system has been deployed.
* @return True if deployed.
*/
bool isDeployed() const;
/**
* @brief Gets the drag coefficient.
* @return Drag coefficient (dimensionless).
*/
double getDragCoefficient() const;
/**
* @brief Gets the reference area.
* @return Reference area (square meters).
*/
double getReferenceArea() const;
/**
* @brief Gets the name of the recovery system.
* @return Recovery system name as a constant reference.
*/
const std::string& getName() const;
private:
std::string name_; ///< Name of the recovery system.
DeploymentType deploymentType_; ///< Deployment trigger type.
double triggerValue_; ///< Trigger value (meters for altitude, seconds for timer).
double dragCoefficient_; ///< Drag coefficient after deployment.
double referenceArea_; ///< Reference area [m2].
bool deployed_; ///< Whether the recovery system has been deployed.
};
#endif // RECOVERYSYSTEM_H

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#ifndef ROCKET_H
#define ROCKET_H
#include <string>
#include <vector>
#include <memory> // for smart pointers
// Forward declarations to keep compile times fast
class Stage;
class FlightState;
class Environment;
/**
* @brief Represents a complete rocket vehicle composed of stages, motors, and recovery systems.
*
* The Rocket class encapsulates the full vehicle configuration, including mass properties,
* center of gravity (CG), center of pressure (CP), and stability margin calculations.
* It is the central object for design, simulation, and flight state management.
*/
class Rocket {
public:
/**
* @brief Constructs a new Rocket with a given name.
* @param name The name of the rocket.
*/
Rocket(const std::string& name);
/**
* @brief Default destructor.
*/
~Rocket() = default;
/**
* @brief Returns the total mass of the rocket (including all stages and payloads).
* @return Total mass in kilograms.
*/
double getTotalMass() const;
/**
* @brief Returns the total remaining propellant mass of all motors.
* @return Propellant mass in kilograms.
*/
double getTotalPropellantMass() const;
/**
* @brief Returns the current center of gravity (CG) of the rocket.
* @return Distance from reference point (e.g., nose tip) in meters.
*/
double getCenterOfGravity() const;
/**
* @brief Returns the current center of pressure (CP) of the rocket.
* @return Distance from reference point (e.g., nose tip) in meters.
*/
double getCenterOfPressure() const;
/**
* @brief Calculates the rocket's stability margin.
*
* Defined as the normalized distance between CG and CP.
* Positive margin indicates stable configuration.
*
* @return Stability margin (calibers).
*/
double getStabilityMargin() const;
/**
* @brief Adds a new stage to the rocket.
* @param stage A unique_ptr to the Stage object to add.
*/
void addStage(std::unique_ptr<Stage> stage);
/**
* @brief Returns a const reference to the list of stages.
* @return Vector of unique_ptr to Stage objects.
*/
const std::vector<std::unique_ptr<Stage>>& getStages() const;
/**
* @brief Recalculates total mass, CG, CP, and stability margin.
*
* Should be called after any modification to the rocket structure.
*/
void updateMassProperties();
/**
* @brief Prepares the rocket for flight simulation.
*
* Typically called once at simulation setup to adjust for environmental conditions.
*
* @param env The environment (atmospheric conditions) at launch.
*/
void prepareForFlight(const Environment& env);
/**
* @brief Applies a given flight state to the rocket.
*
* Intended for advanced simulation stages, including 6-DoF dynamics.
*
* @param state The current flight state.
*/
void applyFlightState(const FlightState& state);
/**
* @brief Gets the name of the rocket.
* @return Rocket name as a constant reference.
*/
const std::string& getName() const;
/**
* @brief Sets the name of the rocket.
* @param name The new name.
*/
void setName(const std::string& name);
private:
std::string name_; ///< Name of the rocket.
std::vector<std::unique_ptr<Stage>> stages_; ///< List of rocket stages.
// Cached mass properties (updated by updateMassProperties())
double totalMass_; ///< Total mass of the rocket [kg].
double totalPropellantMass_; ///< Total remaining propellant mass [kg].
double centerOfGravity_; ///< Center of gravity location [m].
double centerOfPressure_; ///< Center of pressure location [m].
};
#endif // ROCKET_H

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#ifndef STAGE_H
#define STAGE_H
#include <string>
#include <vector>
#include <memory>
class Motor;
class Airframe;
class FinSet;
class RecoverySystem;
class Environment;
class FlightState;
/**
* @brief Represents a single stage of a (potentially) multi-stage rocket.
*
* A Stage contains structural components, motors, recovery systems,
* and manages its own mass properties and separation events.
*/
class Stage {
public:
/**
* @brief Constructs a new Stage with a given name.
* @param name The name of the stage.
*/
Stage(const std::string& name);
/**
* @brief Default destructor.
*/
~Stage() = default;
/**
* @brief Adds a motor to the stage.
* @param motor A unique_ptr to the Motor object.
*/
void addMotor(std::unique_ptr<Motor> motor);
/**
* @brief Sets the airframe for the stage.
* @param airframe A unique_ptr to the Airframe object.
*/
void setAirframe(std::unique_ptr<Airframe> airframe);
/**
* @brief Sets the fin set for the stage.
* @param finSet A unique_ptr to the FinSet object.
*/
void setFinSet(std::unique_ptr<FinSet> finSet);
/**
* @brief Sets the recovery system for the stage.
* @param recovery A unique_ptr to the RecoverySystem object.
*/
void setRecoverySystem(std::unique_ptr<RecoverySystem> recovery);
/**
* @brief Returns the total mass of the stage (structure + motors + payload).
* @return Total mass in kilograms.
*/
double getTotalMass() const;
/**
* @brief Returns the current total propellant mass of all motors in the stage.
* @return Propellant mass in kilograms.
*/
double getTotalPropellantMass() const;
/**
* @brief Updates mass properties (e.g., after motor burn or separation).
*/
void updateMassProperties();
/**
* @brief Prepares the stage for flight simulation (e.g., motor ignition sequencing).
* @param env Launch environment.
*/
void prepareForFlight(const Environment& env);
/**
* @brief Updates the stage based on current flight conditions.
* @param state Current flight state.
*/
void applyFlightState(const FlightState& state);
/**
* @brief Checks if the stage should separate (e.g., after burnout or trigger).
* @return True if ready to separate.
*/
bool checkSeparationEvent() const;
/**
* @brief Checks if the recovery system should deploy (e.g., apogee, velocity triggers).
* @return True if recovery deployment condition met.
*/
bool checkRecoveryEvent() const;
/**
* @brief Gets the length of the airframe.
* @return Length in meters (0 if no airframe assigned).
*/
double getAirframeLength() const;
/**
* @brief Calculates the normal force coefficient contribution from the stage's fins.
* @return Normal force coefficient (dimensionless).
*/
double calculateNormalForceCoefficient() const;
/**
* @brief Gets the name of the stage.
* @return Stage name as a constant reference.
*/
const std::string& getName() const;
/**
* @brief Calculates the total thrust vector produced by all motors in this stage.
* @return Thrust vector (Newton) in body-fixed frame (Z-forward).
*/
std::array<double, 3> getTotalThrust() const;
// Only for testing access!
const std::vector<std::unique_ptr<Motor>>& getMotorsForTesting() const { return motors_; }
private:
std::string name_; ///< Name of the stage.
// Core Components
std::vector<std::unique_ptr<Motor>> motors_; ///< List of motors in this stage.
std::unique_ptr<Airframe> airframe_; ///< Structural body and aerodynamic surfaces.
std::unique_ptr<FinSet> finSet_; ///< Fins for aerodynamic stability.
std::unique_ptr<RecoverySystem> recoverySystem_; ///< Recovery deployment system.
// Cached Mass Properties
double totalMass_; ///< Total mass [kg].
double totalPropellantMass_; ///< Remaining propellant [kg].
// Event Flags
bool separationTriggered_; ///< Flag indicating stage separation event.
bool recoveryDeployed_; ///< Flag indicating recovery system deployment.
};
#endif // STAGE_H

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#ifndef MATPLOTLIBPLOTTER_H
#define MATPLOTLIBPLOTTER_H
#include "Plotter.h"
#include "matplotlibcpp.h"
namespace plt = matplotlibcpp;
/**
* @brief Matplotlib-cpp implementation of the Plotter interface.
*/
class MatplotlibPlotter : public Plotter {
public:
virtual ~MatplotlibPlotter() = default;
void plot2D(const std::vector<double>& xData,
const std::vector<double>& yData,
const std::string& title,
const std::string& xLabel,
const std::string& yLabel) override {
plt::figure();
plt::plot(xData, yData);
// Plot any queued markers
for (const auto& marker : markers_) {
plt::axvline(marker.first, 0.0, 1.0, {{"color", "red"}, {"linestyle", "--"}});
plt::text(marker.first, *std::max_element(yData.begin(), yData.end()),
marker.second);
}
plt::title(title);
plt::xlabel(xLabel);
plt::ylabel(yLabel);
plt::grid(true);
plt::show();
// markers_.clear();
}
void addMarker(double xValue, const std::string& label) override {
markers_.emplace_back(xValue, label);
}
void plotPosVelAcc(const std::vector<double>& time,
const std::vector<double>& altitude,
const std::vector<double>& velocity,
const std::vector<double>& acceleration) override {
plt::figure();
plt::subplot(3, 1, 1);
plt::plot(time, altitude);
plt::title("Altitude vs Time");
plt::ylabel("Altitude (m)");
plt::grid(true);
for (const auto& marker : markers_) {
plt::axvline(marker.first, 0.0, 1.0, {{"color", "red"}, {"linestyle", "--"}});
plt::text(marker.first, *std::max_element(altitude.begin(), altitude.end()),
marker.second);
}
plt::subplot(3, 1, 2);
plt::plot(time, velocity);
plt::title("Velocity vs Time");
plt::ylabel("Velocity (m/s)");
plt::grid(true);
for (const auto& marker : markers_) {
plt::axvline(marker.first, 0.0, 1.0, {{"color", "red"}, {"linestyle", "--"}});
plt::text(marker.first, *std::max_element(velocity.begin(), velocity.end()),
marker.second);
}
plt::subplot(3, 1, 3);
plt::plot(time, acceleration);
plt::title("Acceleration vs Time");
plt::xlabel("Time (s)");
plt::ylabel("Acceleration (m/s2)");
plt::grid(true);
for (const auto& marker : markers_) {
plt::axvline(marker.first, 0.0, 1.0, {{"color", "red"}, {"linestyle", "--"}});
plt::text(marker.first, *std::max_element(acceleration.begin(), acceleration.end()),
marker.second);
}
plt::tight_layout();
plt::show();
// markers_.clear();
}
private:
std::vector<std::pair<double, std::string>> markers_;
};
#endif // MATPLOTLIBPLOTTER_H

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#ifndef PLOTTER_H
#define PLOTTER_H
#include <vector>
#include <string>
/**
* @brief Abstract interface for plotting flight data.
*/
class Plotter {
public:
virtual ~Plotter() = default;
/**
* @brief Plot a 2D line graph.
*
* @param xData Vector of x-axis values (e.g., time)
* @param yData Vector of y-axis values (e.g., altitude)
* @param title Plot title
* @param xLabel Label for x-axis
* @param yLabel Label for y-axis
*/
virtual void plot2D(const std::vector<double>& xData,
const std::vector<double>& yData,
const std::string& title,
const std::string& xLabel,
const std::string& yLabel) = 0;
/**
* @brief Add a vertical marker at a specific x-axis value.
*
* @param xValue The x position (e.g., time in seconds)
* @param label Text label to display near the marker
*/
virtual void addMarker(double xValue, const std::string& label) = 0;
/**
* @brief Plot altitude, velocity, and acceleration together on a single canvas.
*
* @param time Time points
* @param altitude Altitude (Z position)
* @param velocity Vertical velocity
* @param acceleration Vertical acceleration
*/
virtual void plotPosVelAcc(const std::vector<double>& time,
const std::vector<double>& altitude,
const std::vector<double>& velocity,
const std::vector<double>& acceleration) = 0;
};
#endif // PLOTTER_H

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#include "Rocket.h"
int main(int argc, char* argv[]) {
Rocket myRocket("Test Rocket");
}

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// src/Airframe.cpp
#include "Airframe.h"
// Constructor
Airframe::Airframe(const std::string& name,
double length,
double diameter,
double dryMass,
double dragCoefficient)
: name_(name),
length_(length),
diameter_(diameter),
dryMass_(dryMass),
dragCoefficient_(dragCoefficient)
{}
// Public Methods
double Airframe::getLength() const {
return length_;
}
double Airframe::getDiameter() const {
return diameter_;
}
double Airframe::getDryMass() const {
return dryMass_;
}
double Airframe::getDragCoefficient() const {
return dragCoefficient_;
}
double Airframe::getReferenceArea() const {
// Reference area for drag force = ? * (d/2)^2
const double radius = diameter_ / 2.0;
return 3.14159265358979323846 * radius * radius;
}
const std::string& Airframe::getName() const {
return name_;
}

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// src/Environment.cpp
#include "Environment.h"
#include <cmath> // For pow()
namespace {
// Constants
constexpr double R_EARTH = 6371000.0; // Earth radius (meters)
constexpr double G0 = 9.80665; // Gravity at sea level (m/s2)
constexpr double R_AIR = 287.05; // Specific gas constant for air (J/kg/K)
// US Standard Atmosphere 1976 layers up to 86 km
struct AtmosphereLayer {
double baseAltitude; // meters
double baseTemperature; // Kelvin
double basePressure; // Pascals
double lapseRate; // K/m
};
// Base layer properties precomputed
static const AtmosphereLayer AtmosphereLayers[] = {
{0.0, 288.15, 101325.0, -0.0065},
{11000.0, 216.65, 22632.06, 0.0},
{20000.0, 216.65, 5474.889, 0.001},
{32000.0, 228.65, 868.019, 0.0028},
{47000.0, 270.65, 110.906, 0.0},
{51000.0, 270.65, 66.93887, -0.0028},
{71000.0, 214.65, 3.956420, -0.002},
};
static constexpr int NumAtmosphereLayers = sizeof(AtmosphereLayers) / sizeof(AtmosphereLayer);
//const AtmosphereLayer* Environment::findAtmosphereLayer(double altitude) const {
const AtmosphereLayer* findAtmosphereLayer(double altitude) {
for (int i = NumAtmosphereLayers - 1; i >= 0; --i) {
if (altitude >= AtmosphereLayers[i].baseAltitude) {
return &AtmosphereLayers[i];
}
}
// Should never happen, but fallback
return &AtmosphereLayers[0];
}
} // namespace
// Constructor
Environment::Environment(double surfacePressure,
double surfaceTemperature,
double launchAltitude)
: surfacePressure_(surfacePressure),
surfaceTemperature_(surfaceTemperature),
launchAltitude_(launchAltitude)
{}
// Public Methods
double Environment::getAirDensity(double altitude) const {
double temp = getStandardTemperature(altitude);
double pressure = getStandardPressure(altitude);
return pressure / (R_AIR * temp);
}
double Environment::getGravity(double altitude) const {
return G0 * std::pow(R_EARTH / (R_EARTH + altitude), 2);
}
double Environment::getSurfacePressure() const {
return surfacePressure_;
}
double Environment::getSurfaceTemperature() const {
return surfaceTemperature_;
}
double Environment::getLaunchAltitude() const {
return launchAltitude_;
}
// Private Methods
double Environment::getStandardTemperature(double altitude) const {
const AtmosphereLayer* layer = findAtmosphereLayer(altitude);
double deltaH = altitude - layer->baseAltitude;
return layer->baseTemperature + layer->lapseRate * deltaH;
}
double Environment::getStandardPressure(double altitude) const {
const AtmosphereLayer* layer = findAtmosphereLayer(altitude);
double deltaH = altitude - layer->baseAltitude;
double T = getStandardTemperature(altitude);
double T0 = layer->baseTemperature;
double P0 = layer->basePressure;
double a = layer->lapseRate;
if (a == 0.0) {
// Isothermal layer
return P0 * std::exp(-G0 * deltaH / (R_AIR * T0));
} else {
// Gradient layer
return P0 * std::pow(T0 / T, (G0 / (a * R_AIR)));
}
}

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// src/FinSet.cpp
#include "FinSet.h"
// Constructor
FinSet::FinSet(const std::string& name,
int numberOfFins,
double rootChord,
double tipChord,
double span,
double sweepLength,
double thickness,
double materialDensity)
: name_(name),
numberOfFins_(numberOfFins),
rootChord_(rootChord),
tipChord_(tipChord),
span_(span),
sweepLength_(sweepLength),
thickness_(thickness),
materialDensity_(materialDensity)
{}
// Public Methods
double FinSet::calculateNormalForceCoefficient() const {
// Simplified thin airfoil theory estimation
// Cn_alpha ? (number of fins) * (planform area) / (reference area)
// Assume reference area (body) is much smaller for initial versions
double planformArea = calculateSingleFinArea();
double effectiveArea = numberOfFins_ * planformArea;
// Return some typical small value scaled by fin count and area
// (In reality, depends heavily on Mach number and sweep angle)
return 2.0 * effectiveArea; // Very basic approximation
}
double FinSet::calculateCenterOfPressure() const {
// Approximate aerodynamic center (x_cp) for trapezoidal fin
// Formula from Barrowman equations (simplified for low-speed flight)
double meanAerodynamicChord = (rootChord_ + tipChord_) / 2.0;
double x_leading_edge = sweepLength_; // Assume sweepLength at root
double x_cp = x_leading_edge + (meanAerodynamicChord / 2.0);
return x_cp;
}
double FinSet::calculateDragArea() const {
// Drag Area = Cd_fins * total projected area
double Cd_fin = 0.01; // Approximate for small rocket fins
double planformArea = calculateSingleFinArea();
double totalArea = numberOfFins_ * planformArea;
return Cd_fin * totalArea;
}
double FinSet::calculateMass() const {
// Volume = area * thickness
double singleArea = calculateSingleFinArea();
double volume = singleArea * thickness_;
double mass = numberOfFins_ * volume * materialDensity_;
return mass;
}
const std::string& FinSet::getName() const {
return name_;
}
// Private Methods
double FinSet::calculateSingleFinArea() const {
// Area of a trapezoid: A = 0.5 * (root + tip) * span
return 0.5 * (rootChord_ + tipChord_) * span_;
}

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// src/FlightSimulator.cpp
#include "FlightSimulator.h"
#include "Rocket.h"
#include "Environment.h"
#include "ForcesModel.h"
#include "FlightState.h"
#include "Integrator.h"
#include "Stage.h"
#include "Motor.h"
#include <iostream> // Optional for debug output
// Constructor
FlightSimulator::FlightSimulator(std::shared_ptr<Rocket> rocket,
std::shared_ptr<Environment> environment)
: rocket_(rocket),
environment_(environment),
forcesModel_(std::make_unique<ForcesModel>(rocket, environment)),
integrator_(std::make_unique<Integrator>(rocket, forcesModel_)),
hasLaunched_(false) // <-- initialize launch flag
{}
// Public Methods
void FlightSimulator::run(double maxSimulationTime, double timeStep) {
initialize();
FlightState state;
state.setTime(0.0);
// Initial position: launch pad
state.setPosition({0.0, 0.0, environment_->getLaunchAltitude()});
state.setVelocity({0.0, 0.0, 0.0});
state.setAcceleration({0.0, 0.0, 0.0});
flightLog_.clear();
flightLog_.push_back(state);
double currentTime = 0.0;
while (currentTime < maxSimulationTime) {
step(timeStep);
const auto& latestState = flightLog_.back();
if (checkTermination(latestState)) {
break;
}
currentTime += timeStep;
}
}
const std::vector<FlightState>& FlightSimulator::getFlightLog() const {
return flightLog_;
}
// Private Methods
void FlightSimulator::initialize() {
rocket_->prepareForFlight(*environment_);
}
void FlightSimulator::step(double deltaTime) {
FlightState currentState = flightLog_.back(); // Copy latest state
updateMotors(deltaTime); // <-- motors must advance ignition timers
// Compute forces and integrate motion
auto netForce = forcesModel_->computeNetForce(currentState);
integrator_->step(currentState, netForce, deltaTime);
rocket_->applyFlightState(currentState);
handleEvents(currentState);
currentState.setTime(currentState.getTime() + deltaTime);
flightLog_.push_back(currentState);
}
void FlightSimulator::updateMotors(double deltaTime) {
for (const auto& stage : rocket_->getStages()) {
for (auto& motor : stage->getMotorsForTesting()) {
motor->update(deltaTime);
}
}
}
void FlightSimulator::handleEvents(FlightState& /*state*/) {
// TODO: Extend this to handle recovery deployment and staging events.
}
bool FlightSimulator::checkTermination(const FlightState& state) {
const double altitude = state.getPosition()[2];
const double verticalVelocity = state.getVelocity()[2];
if (!hasLaunched_) {
// Detect real liftoff: significant velocity or altitude gain
if (verticalVelocity > 5.0 || altitude > environment_->getLaunchAltitude() + 5.0) {
hasLaunched_ = true;
}
return false; // Always continue until launched
}
// After launch, normal landing detection
if (altitude <= environment_->getLaunchAltitude() + 1.0 && verticalVelocity <= 1.0) {
return true; // Landed
}
return false; // Still flying
}

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// src/FlightState.cpp
#include "FlightState.h"
// Constructor
FlightState::FlightState()
: position_{0.0, 0.0, 0.0},
velocity_{0.0, 0.0, 0.0},
acceleration_{0.0, 0.0, 0.0},
orientation_{1.0, 0.0, 0.0, 0.0}, // Unit quaternion by default
angularVelocity_{0.0, 0.0, 0.0},
time_(0.0)
{}
// Public Methods
const std::array<double, 3>& FlightState::getPosition() const {
return position_;
}
void FlightState::setPosition(const std::array<double, 3>& pos) {
position_ = pos;
}
const std::array<double, 3>& FlightState::getVelocity() const {
return velocity_;
}
void FlightState::setVelocity(const std::array<double, 3>& vel) {
velocity_ = vel;
}
const std::array<double, 3>& FlightState::getAcceleration() const {
return acceleration_;
}
void FlightState::setAcceleration(const std::array<double, 3>& acc) {
acceleration_ = acc;
}
const std::array<double, 4>& FlightState::getOrientation() const {
return orientation_;
}
void FlightState::setOrientation(const std::array<double, 4>& quat) {
orientation_ = quat;
}
const std::array<double, 3>& FlightState::getAngularVelocity() const {
return angularVelocity_;
}
void FlightState::setAngularVelocity(const std::array<double, 3>& angVel) {
angularVelocity_ = angVel;
}
double FlightState::getTime() const {
return time_;
}
void FlightState::setTime(double time) {
time_ = time;
}

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// src/ForcesModel.cpp
#include "ForcesModel.h"
#include "Rocket.h"
#include "Stage.h"
#include "Motor.h"
#include "Airframe.h"
#include "FinSet.h"
#include "RecoverySystem.h"
#include "Environment.h"
#include "FlightState.h"
#include <cmath> // For sqrt
// Constructor
ForcesModel::ForcesModel(std::shared_ptr<Rocket> rocket,
std::shared_ptr<Environment> environment)
: rocket_(rocket),
environment_(environment)
{}
// Public Methods
std::array<double, 3> ForcesModel::computeNetForce(const FlightState& state) const {
std::array<double, 3> netForce = {0.0, 0.0, 0.0};
// Thrust
auto thrust = computeThrust(state);
for (int i = 0; i < 3; ++i) {
netForce[i] += thrust[i];
}
// Aerodynamic Drag
auto drag = computeAerodynamicDrag(state);
for (int i = 0; i < 3; ++i) {
netForce[i] += drag[i];
}
// Gravity
auto gravity = computeGravity(state);
for (int i = 0; i < 3; ++i) {
netForce[i] += gravity[i];
}
return netForce;
}
std::array<double, 3> ForcesModel::computeNetMoment(const FlightState& /*state*/) const {
// Placeholder: No moments in basic 3-DoF model
return {0.0, 0.0, 0.0};
}
// Private Methods
std::array<double, 3> ForcesModel::computeAerodynamicDrag(const FlightState& state) const {
std::array<double, 3> dragForce = {0.0, 0.0, 0.0};
// Get air density at altitude
double altitude = state.getPosition()[2];
double airDensity = environment_->getAirDensity(altitude);
// Get velocity
const auto& velocity = state.getVelocity();
double speed = std::sqrt(velocity[0]*velocity[0] +
velocity[1]*velocity[1] +
velocity[2]*velocity[2]);
if (speed == 0.0) {
return dragForce; // No drag if stationary
}
// Drag direction is opposite to velocity
std::array<double, 3> dragDirection = {-velocity[0]/speed,
-velocity[1]/speed,
-velocity[2]/speed};
// Approximate total reference area and Cd
double totalReferenceArea = 0.0;
double effectiveDragCoefficient = 0.0;
for (const auto& stage : rocket_->getStages()) {
if (stage->getAirframeLength() > 0.0) {
totalReferenceArea += stage->getAirframeLength(); // Simple approx
effectiveDragCoefficient += 0.75; // Assume average Cd
}
}
// Recovery drag if deployed
for (const auto& stage : rocket_->getStages()) {
// Assume only one recovery system per stage for now
// (Can easily extend)
auto recovery = stage->checkRecoveryEvent(); // Would return a bool now
if (recovery) {
totalReferenceArea += 1.0; // Large area
effectiveDragCoefficient += 1.5; // Large Cd
}
}
if (totalReferenceArea <= 0.0) {
totalReferenceArea = 0.01; // Prevent division by zero
}
// Compute drag force magnitude
double dragMagnitude = 0.5 * airDensity * speed * speed *
effectiveDragCoefficient * totalReferenceArea;
// Apply direction
for (int i = 0; i < 3; ++i) {
dragForce[i] = dragMagnitude * dragDirection[i];
}
return dragForce;
}
std::array<double, 3> ForcesModel::computeThrust(const FlightState& /*state*/) const {
std::array<double, 3> thrustForce = {0.0, 0.0, 0.0};
for (const auto& stage : rocket_->getStages()) {
auto stageThrust = stage->getTotalThrust();
for (int i = 0; i < 3; ++i) {
thrustForce[i] += stageThrust[i];
}
}
return thrustForce;
}
std::array<double, 3> ForcesModel::computeGravity(const FlightState& state) const {
std::array<double, 3> gravityForce = {0.0, 0.0, 0.0};
double altitude = state.getPosition()[2];
// TODO: Make this more robust. Don't accelerate through the ground
if(altitude <= 0.0) {
return {0.0, 0.0, 0.0};
}
double g = environment_->getGravity(altitude);
double totalMass = rocket_->getTotalMass();
gravityForce[2] = -totalMass * g; // Gravity acts downward
return gravityForce;
}

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// src/Integrator.cpp
#include "Integrator.h"
#include "Rocket.h"
#include "ForcesModel.h"
#include "FlightState.h"
#include <array>
// Constructor
Integrator::Integrator(std::shared_ptr<Rocket> rocket,
std::shared_ptr<ForcesModel> forcesModel)
: rocket_(rocket),
forcesModel_(forcesModel)
{}
// Public Methods
void Integrator::step(FlightState& state, const std::array<double, 3>& netForce, double deltaTime) {
// Basic Euler integration for now
eulerIntegration(state, netForce, deltaTime);
}
// Private Methods
void Integrator::eulerIntegration(FlightState& state,
const std::array<double, 3>& netForce,
double deltaTime) {
// Get current values
auto position = state.getPosition();
auto velocity = state.getVelocity();
auto acceleration = state.getAcceleration();
double mass = rocket_->getTotalMass();
// Update acceleration: a = F / m
if (mass > 0.0) {
for (int i = 0; i < 3; ++i) {
acceleration[i] = netForce[i] / mass;
}
} else {
acceleration = {0.0, 0.0, 0.0}; // Safety fallback
}
// Update velocity: v = v + a * dt
for (int i = 0; i < 3; ++i) {
velocity[i] += acceleration[i] * deltaTime;
}
// Update position: x = x + v * dt
for (int i = 0; i < 3; ++i) {
position[i] += velocity[i] * deltaTime;
}
// Update state
state.setAcceleration(acceleration);
state.setVelocity(velocity);
state.setPosition(position);
}

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#include "Motor.h"
#include <algorithm> // for std::lower_bound
// Constructor
Motor::Motor(const std::string& name, double initialPropellantMass, double totalImpulse)
: name_(name),
initialPropellantMass_(initialPropellantMass),
remainingPropellantMass_(initialPropellantMass),
totalImpulse_(totalImpulse),
ignitionTime_(0.0),
ignited_(false),
burnedOut_(false),
currentThrust_(0.0)
{}
// Public Methods
void Motor::addThrustDataPoint(double time, double thrust) {
thrustCurve_.emplace_back(time, thrust);
}
void Motor::ignite() {
ignitionTime_ = 0.0;
ignited_ = true;
burnedOut_ = false;
}
void Motor::update(double deltaTime) {
if (!ignited_ || burnedOut_) {
currentThrust_ = 0.0;
return;
}
ignitionTime_ += deltaTime;
// Check if we exceed thrust curve duration
if (!thrustCurve_.empty() && ignitionTime_ > thrustCurve_.back().first) {
burnedOut_ = true;
currentThrust_ = 0.0;
remainingPropellantMass_ = 0.0;
return;
}
// Update current thrust using interpolation
currentThrust_ = interpolateThrust(ignitionTime_);
// Estimate propellant burn proportional to impulse usage
double impulsePerSecond = totalImpulse_ / (thrustCurve_.back().first); // approximate
double propellantConsumptionRate = initialPropellantMass_ / (totalImpulse_ / impulsePerSecond);
double deltaPropellant = propellantConsumptionRate * deltaTime;
remainingPropellantMass_ = std::max(remainingPropellantMass_ - deltaPropellant, 0.0);
}
double Motor::getCurrentThrust() const {
return currentThrust_;
}
bool Motor::isBurnedOut() const {
return burnedOut_;
}
double Motor::getRemainingPropellantMass() const {
return remainingPropellantMass_;
}
const std::string& Motor::getName() const {
return name_;
}
// Private Methods
double Motor::interpolateThrust(double time) const {
if (thrustCurve_.empty()) {
return 0.0;
}
// If time is outside bounds
if (time <= thrustCurve_.front().first) {
return thrustCurve_.front().second;
}
if (time >= thrustCurve_.back().first) {
return thrustCurve_.back().second;
}
// Find the two points around the current time
auto it = std::lower_bound(thrustCurve_.begin(), thrustCurve_.end(), std::make_pair(time, 0.0),
[](const std::pair<double, double>& a, const std::pair<double, double>& b) {
return a.first < b.first;
});
if (it == thrustCurve_.begin()) {
return it->second;
}
auto itPrev = std::prev(it);
double t0 = itPrev->first;
double t1 = it->first;
double f0 = itPrev->second;
double f1 = it->second;
// Linear interpolation
double thrust = f0 + (f1 - f0) * (time - t0) / (t1 - t0);
return thrust;
}

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// src/RecoverySystem.cpp
#include "RecoverySystem.h"
// Constructor
RecoverySystem::RecoverySystem(const std::string& name,
DeploymentType deploymentType,
double triggerValue,
double dragCoefficient,
double referenceArea)
: name_(name),
deploymentType_(deploymentType),
triggerValue_(triggerValue),
dragCoefficient_(dragCoefficient),
referenceArea_(referenceArea),
deployed_(false)
{}
// Public Methods
bool RecoverySystem::checkDeploymentCondition(double altitude,
double velocity,
double time,
bool atApogee) const {
if (deployed_) {
return false; // Already deployed
}
switch (deploymentType_) {
case DeploymentType::Apogee:
return atApogee;
case DeploymentType::Altitude:
return (altitude <= triggerValue_);
case DeploymentType::Timer:
return (time >= triggerValue_);
default:
return false;
}
}
void RecoverySystem::deploy() {
deployed_ = true;
}
bool RecoverySystem::isDeployed() const {
return deployed_;
}
double RecoverySystem::getDragCoefficient() const {
return dragCoefficient_;
}
double RecoverySystem::getReferenceArea() const {
return referenceArea_;
}
const std::string& RecoverySystem::getName() const {
return name_;
}

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#include "Rocket.h"
#include "Stage.h"
#include "Environment.h"
#include "FlightState.h"
#include "Airframe.h"
#include "FinSet.h"
#include "RecoverySystem.h"
#include "Motor.h"
#include <numeric> // for std::accumulate
// Constructor
Rocket::Rocket(const std::string& name)
: name_(name),
totalMass_(0.0),
totalPropellantMass_(0.0),
centerOfGravity_(0.0),
centerOfPressure_(0.0)
{}
// Public Methods
double Rocket::getTotalMass() const {
return totalMass_;
}
double Rocket::getTotalPropellantMass() const {
return totalPropellantMass_;
}
double Rocket::getCenterOfGravity() const {
return centerOfGravity_;
}
double Rocket::getCenterOfPressure() const {
return centerOfPressure_;
}
double Rocket::getStabilityMargin() const {
// Positive margin (CG ahead of CP) indicates stability.
// Normalize by diameter or calibers if needed later.
return centerOfGravity_ - centerOfPressure_;
}
void Rocket::addStage(std::unique_ptr<Stage> stage) {
stages_.emplace_back(std::move(stage));
}
const std::vector<std::unique_ptr<Stage>>& Rocket::getStages() const {
return stages_;
}
void Rocket::updateMassProperties() {
totalMass_ = 0.0;
totalPropellantMass_ = 0.0;
centerOfGravity_ = 0.0;
centerOfPressure_ = 0.0;
if (stages_.empty()) {
return;
}
double weightedCGSum = 0.0;
double weightedCPSum = 0.0;
double currentZ = 0.0; // Assume stacking stages vertically
for (const auto& stage : stages_) {
stage->updateMassProperties();
double stageMass = stage->getTotalMass();
double stagePropellantMass = stage->getTotalPropellantMass();
totalMass_ += stageMass;
totalPropellantMass_ += stagePropellantMass;
// For center of gravity, weight by mass
weightedCGSum += (currentZ + stage->getAirframeLength() / 2.0) * stageMass;
// For center of pressure, weight by aerodynamic properties
weightedCPSum += (currentZ + stage->getAirframeLength() / 2.0) * stage->calculateNormalForceCoefficient();
currentZ += stage->getAirframeLength(); // Move up for next stage
}
centerOfGravity_ = weightedCGSum / totalMass_;
// Note: CP calculation here is a rough average based on fin normal forces
centerOfPressure_ = weightedCPSum /
std::accumulate(stages_.begin(), stages_.end(), 0.0,
[](double sum, const std::unique_ptr<Stage>& s) {
return sum + s->calculateNormalForceCoefficient();
});
}
void Rocket::prepareForFlight(const Environment& env) {
for (auto& stage : stages_) {
stage->prepareForFlight(env);
}
}
void Rocket::applyFlightState(const FlightState& state) {
for (auto& stage : stages_) {
stage->applyFlightState(state);
}
}
const std::string& Rocket::getName() const {
return name_;
}
void Rocket::setName(const std::string& name) {
name_ = name;
}

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// src/Stage.cpp
#include "Stage.h"
#include "Motor.h"
#include "Airframe.h"
#include "FinSet.h"
#include "RecoverySystem.h"
#include "Environment.h"
#include "FlightState.h"
#include <numeric> // for std::accumulate
#include <algorithm> // for std::all_of
// Constructor
Stage::Stage(const std::string& name)
: name_(name),
totalMass_(0.0),
totalPropellantMass_(0.0),
separationTriggered_(false),
recoveryDeployed_(false)
{}
// Public Methods
void Stage::addMotor(std::unique_ptr<Motor> motor) {
motors_.emplace_back(std::move(motor));
}
void Stage::setAirframe(std::unique_ptr<Airframe> airframe) {
airframe_ = std::move(airframe);
}
void Stage::setFinSet(std::unique_ptr<FinSet> finSet) {
finSet_ = std::move(finSet);
}
void Stage::setRecoverySystem(std::unique_ptr<RecoverySystem> recovery) {
recoverySystem_ = std::move(recovery);
}
double Stage::getTotalMass() const {
return totalMass_;
}
double Stage::getTotalPropellantMass() const {
return totalPropellantMass_;
}
void Stage::updateMassProperties() {
totalMass_ = 0.0;
totalPropellantMass_ = 0.0;
// Add airframe dry mass
if (airframe_) {
totalMass_ += airframe_->getDryMass();
}
// Add motor masses
for (const auto& motor : motors_) {
totalMass_ += motor->getRemainingPropellantMass();
totalPropellantMass_ += motor->getRemainingPropellantMass();
}
// Add fin mass
if (finSet_) {
totalMass_ += finSet_->calculateMass();
}
}
void Stage::prepareForFlight(const Environment& env) {
// Ignite all motors immediately at start (for simplicity now)
for (auto& motor : motors_) {
motor->ignite();
}
}
void Stage::applyFlightState(const FlightState& state) {
// Placeholder for future advanced logic (flex, vibration)
// Currently nothing needed at 3-DoF level
}
bool Stage::checkSeparationEvent() const {
// Simple rule: if all motors burned out, stage is done
bool allBurnedOut = std::all_of(motors_.begin(), motors_.end(),
[](const std::unique_ptr<Motor>& motor) {
return motor->isBurnedOut();
});
return allBurnedOut;
}
bool Stage::checkRecoveryEvent() const {
if (!recoverySystem_ || recoverySystem_->isDeployed()) {
return false;
}
// RecoverySystem handles checking altitude, velocity, etc.
// Here we just provide example usage
return false; // Actual deployment logic will be done in FlightSimulator::handleEvents()
}
const std::string& Stage::getName() const {
return name_;
}
double Stage::getAirframeLength() const {
if (airframe_) {
return airframe_->getLength();
}
return 0.0;
}
double Stage::calculateNormalForceCoefficient() const {
if (finSet_) {
return finSet_->calculateNormalForceCoefficient();
}
return 0.0;
}
std::array<double, 3> Stage::getTotalThrust() const {
std::array<double, 3> totalThrust = {0.0, 0.0, 0.0};
for (const auto& motor : motors_) {
// For now, assume thrust always points along +Z axis
totalThrust[2] += motor->getCurrentThrust();
}
return totalThrust;
}

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// tests/test_airframe.cpp
#include "vendor/catch_amalgamated.hpp"
#include "Airframe.h"
// Bring Approx into scope
using Catch::Approx;
TEST_CASE("Airframe basic behavior", "[airframe]") {
Airframe airframe(
"TestBody",
1.5, // length meters
0.1, // diameter meters
2.0, // dry mass kg
0.75 // drag coefficient (unitless)
);
SECTION("Airframe properties are correctly initialized") {
REQUIRE(airframe.getLength() == Approx(1.5));
REQUIRE(airframe.getDiameter() == Approx(0.1));
REQUIRE(airframe.getDryMass() == Approx(2.0));
REQUIRE(airframe.getDragCoefficient() == Approx(0.75));
REQUIRE(airframe.getName() == "TestBody");
}
SECTION("Reference area calculation is correct") {
double expectedArea = 3.14159265358979323846 * (0.1/2.0) * (0.1/2.0);
REQUIRE(airframe.getReferenceArea() == Approx(expectedArea).epsilon(0.001));
}
}

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// tests/test_finset.cpp
#include "vendor/catch_amalgamated.hpp"
#include "FinSet.h"
// Bring Approx into scope
using Catch::Approx;
TEST_CASE("FinSet basic behavior", "[finset]") {
FinSet finSet(
"TestFins",
3, // Number of fins
0.15, // Root chord (m)
0.10, // Tip chord (m)
0.10, // Span (m)
0.05, // Sweep length (m)
0.003, // Thickness (m)
500.0 // Material density (kg/m3)
);
SECTION("FinSet properties are initialized correctly") {
REQUIRE(finSet.getName() == "TestFins");
}
SECTION("Normal force coefficient is positive") {
double Cn = finSet.calculateNormalForceCoefficient();
REQUIRE(Cn > 0.0);
}
SECTION("Center of pressure is reasonable") {
double cp = finSet.calculateCenterOfPressure();
REQUIRE(cp > 0.0);
}
SECTION("Drag area calculation is reasonable") {
double dragArea = finSet.calculateDragArea();
REQUIRE(dragArea > 0.0);
}
SECTION("Mass calculation is reasonable") {
double mass = finSet.calculateMass();
REQUIRE(mass > 0.0);
}
}

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// tests/test_flightstate.cpp
#include "vendor/catch_amalgamated.hpp"
#include "FlightState.h"
// Bring Approx into scope
using Catch::Approx;
TEST_CASE("FlightState basic behavior", "[flightstate]") {
FlightState state;
SECTION("Initial state values are zero or identity") {
REQUIRE(state.getPosition()[0] == Approx(0.0));
REQUIRE(state.getVelocity()[0] == Approx(0.0));
REQUIRE(state.getAcceleration()[0] == Approx(0.0));
REQUIRE(state.getOrientation()[0] == Approx(1.0)); // w component of unit quaternion
REQUIRE(state.getOrientation()[1] == Approx(0.0));
REQUIRE(state.getTime() == Approx(0.0));
}
SECTION("Can set and get position") {
std::array<double, 3> pos = {10.0, 20.0, 30.0};
state.setPosition(pos);
REQUIRE(state.getPosition()[0] == Approx(10.0));
REQUIRE(state.getPosition()[1] == Approx(20.0));
REQUIRE(state.getPosition()[2] == Approx(30.0));
}
SECTION("Can set and get velocity") {
std::array<double, 3> vel = {1.0, 2.0, 3.0};
state.setVelocity(vel);
REQUIRE(state.getVelocity()[0] == Approx(1.0));
REQUIRE(state.getVelocity()[1] == Approx(2.0));
REQUIRE(state.getVelocity()[2] == Approx(3.0));
}
SECTION("Can set and get acceleration") {
std::array<double, 3> acc = {-1.0, -2.0, -3.0};
state.setAcceleration(acc);
REQUIRE(state.getAcceleration()[0] == Approx(-1.0));
REQUIRE(state.getAcceleration()[1] == Approx(-2.0));
REQUIRE(state.getAcceleration()[2] == Approx(-3.0));
}
SECTION("Can set and get orientation") {
std::array<double, 4> quat = {0.707, 0.0, 0.707, 0.0}; // 90? rotation around Y axis
state.setOrientation(quat);
REQUIRE(state.getOrientation()[0] == Approx(0.707));
REQUIRE(state.getOrientation()[2] == Approx(0.707));
}
SECTION("Can set and get angular velocity") {
std::array<double, 3> angVel = {0.01, 0.02, 0.03};
state.setAngularVelocity(angVel);
REQUIRE(state.getAngularVelocity()[0] == Approx(0.01));
REQUIRE(state.getAngularVelocity()[1] == Approx(0.02));
REQUIRE(state.getAngularVelocity()[2] == Approx(0.03));
}
SECTION("Can set and get time") {
state.setTime(5.5);
REQUIRE(state.getTime() == Approx(5.5));
}
}

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#define CATCH_CONFIG_MAIN
#include "vendor/catch_amalgamated.hpp"

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// tests/test_motor.cpp
#include "vendor/catch_amalgamated.hpp"
#include "Motor.h"
// Bring Approx into local scope
using Catch::Approx;
TEST_CASE("Motor basic behavior", "[motor]") {
Motor testMotor("TestMotor", 0.05, 50.0); // 50 grams propellant, 50 Ns impulse
SECTION("Motor should not be ignited at start") {
REQUIRE(testMotor.isBurnedOut() == false);
REQUIRE(testMotor.getCurrentThrust() == Approx(0.0));
}
SECTION("Motor thrust curve points can be added") {
testMotor.addThrustDataPoint(0.0, 0.0);
testMotor.addThrustDataPoint(0.5, 100.0);
testMotor.addThrustDataPoint(1.0, 0.0);
testMotor.ignite();
testMotor.update(0.25); // halfway to 0.5 seconds
// Should interpolate between 0 and 100 N
REQUIRE(testMotor.getCurrentThrust() > 0.0);
REQUIRE(testMotor.isBurnedOut() == false);
}
SECTION("Motor ignition starts thrust production") {
testMotor.addThrustDataPoint(0.0, 0.0);
testMotor.addThrustDataPoint(0.1, 50.0);
testMotor.addThrustDataPoint(1.5, 50.0);
testMotor.addThrustDataPoint(1.6, 0.0);
testMotor.ignite();
testMotor.update(0.1); // 100ms
REQUIRE(testMotor.getCurrentThrust() > 0.0);
REQUIRE(testMotor.isBurnedOut() == false);
}
SECTION("Motor burns out after thrust curve ends") {
testMotor.addThrustDataPoint(0.0, 0.0);
testMotor.addThrustDataPoint(0.5, 50.0);
testMotor.addThrustDataPoint(1.0, 0.0);
testMotor.ignite();
testMotor.update(2.0); // 2 seconds, way past end
REQUIRE(testMotor.isBurnedOut() == true);
REQUIRE(testMotor.getCurrentThrust() == Approx(0.0));
}
SECTION("Remaining propellant decreases over time") {
testMotor.addThrustDataPoint(0.0, 0.0);
testMotor.addThrustDataPoint(0.5, 100.0);
testMotor.addThrustDataPoint(1.0, 0.0);
testMotor.ignite();
double initialMass = testMotor.getRemainingPropellantMass();
testMotor.update(0.5); // Burn for 0.5 seconds
double remainingMass = testMotor.getRemainingPropellantMass();
REQUIRE(remainingMass < initialMass);
REQUIRE(remainingMass >= 0.0);
}
}

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// tests/test_recoverysystem.cpp
#include "vendor/catch_amalgamated.hpp"
#include "RecoverySystem.h"
// Bring Approx into scope
using Catch::Approx;
TEST_CASE("RecoverySystem basic behavior", "[recoverysystem]") {
RecoverySystem apogeeRecovery(
"MainChute",
RecoverySystem::DeploymentType::Apogee,
0.0, // trigger value ignored for Apogee type
1.5, // drag coefficient
2.0 // reference area (m2)
);
RecoverySystem altitudeRecovery(
"DrogueChute",
RecoverySystem::DeploymentType::Altitude,
300.0, // trigger altitude (meters)
1.2, // drag coefficient
1.0 // reference area (m2)
);
RecoverySystem timerRecovery(
"BackupChute",
RecoverySystem::DeploymentType::Timer,
15.0, // trigger time (seconds)
1.8, // drag coefficient
1.5 // reference area (m2)
);
SECTION("RecoverySystem properties are correctly initialized") {
REQUIRE(apogeeRecovery.getName() == "MainChute");
REQUIRE(altitudeRecovery.getDragCoefficient() == Approx(1.2));
REQUIRE(timerRecovery.getReferenceArea() == Approx(1.5));
}
SECTION("Apogee deployment triggers correctly") {
bool deploy = apogeeRecovery.checkDeploymentCondition(
1000.0, // altitude
-10.0, // descending
10.0, // flight time
true // atApogee
);
REQUIRE(deploy == true);
}
SECTION("Altitude deployment triggers correctly") {
bool deployHigh = altitudeRecovery.checkDeploymentCondition(
500.0, // altitude
-10.0, // descending
10.0, // flight time
false // not necessarily at apogee
);
REQUIRE(deployHigh == false);
bool deployLow = altitudeRecovery.checkDeploymentCondition(
250.0, // below trigger altitude
-10.0, // descending
10.0, // flight time
false
);
REQUIRE(deployLow == true);
}
SECTION("Timer deployment triggers correctly") {
bool deployEarly = timerRecovery.checkDeploymentCondition(
500.0, // altitude
-10.0, // descending
10.0, // flight time (too early)
false
);
REQUIRE(deployEarly == false);
bool deployLate = timerRecovery.checkDeploymentCondition(
500.0, // altitude
-10.0, // descending
20.0, // flight time (after timer)
false
);
REQUIRE(deployLate == true);
}
SECTION("Deployment state tracks correctly") {
RecoverySystem testRecovery(
"TestChute",
RecoverySystem::DeploymentType::Apogee,
0.0,
1.5,
2.0
);
REQUIRE(testRecovery.isDeployed() == false);
testRecovery.deploy();
REQUIRE(testRecovery.isDeployed() == true);
// Once deployed, checkDeploymentCondition should return false
REQUIRE(testRecovery.checkDeploymentCondition(
1000.0, -10.0, 10.0, true) == false);
}
}

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// tests/test_rocket_simulator.cpp
#include "vendor/catch_amalgamated.hpp"
#include "Rocket.h"
#include "Stage.h"
#include "Motor.h"
#include "Airframe.h"
#include "FinSet.h"
#include "RecoverySystem.h"
#include "Environment.h"
#include "FlightSimulator.h"
#include "FlightState.h"
#include "Integrator.h"
// Bring Approx into scope
using Catch::Approx;
TEST_CASE("Rocket and FlightSimulator integration test", "[integration]") {
// Setup Environment
auto environment = std::make_shared<Environment>(
101325.0, // surface pressure (Pa)
288.15, // surface temperature (K)
0.0 // launch altitude (m)
);
// Setup Rocket
auto rocket = std::make_shared<Rocket>("IntegrationTestRocket");
auto stage = std::make_unique<Stage>("Stage1");
// Motor
auto motor = std::make_unique<Motor>("TestMotor", 0.05, 50.0); // 50g propellant
motor->addThrustDataPoint(0.0, 0.0);
motor->addThrustDataPoint(0.1, 50.0);
motor->addThrustDataPoint(1.5, 50.0);
motor->addThrustDataPoint(1.6, 0.0);
stage->addMotor(std::move(motor));
// Airframe
auto airframe = std::make_unique<Airframe>("Airframe1",
1.5, // length
0.1, // diameter
2.0, // dry mass
0.75 // drag coefficient
);
stage->setAirframe(std::move(airframe));
// FinSet
auto finset = std::make_unique<FinSet>("TestFins",
3, // 3 fins
0.15, // root chord
0.10, // tip chord
0.10, // span
0.05, // sweep
0.003, // thickness
500.0 // density
);
stage->setFinSet(std::move(finset));
// Recovery
auto recovery = std::make_unique<RecoverySystem>("MainChute",
RecoverySystem::DeploymentType::Apogee,
0.0, // no trigger value needed
1.5, // Cd
2.0 // area
);
stage->setRecoverySystem(std::move(recovery));
rocket->addStage(std::move(stage));
rocket->updateMassProperties();
// Setup FlightSimulator
FlightSimulator simulator(rocket, environment);
SECTION("Full flight simulation runs and reaches apogee") {
simulator.run(20.0, 0.01); // Simulate 20 seconds with 10ms step
const auto& flightLog = simulator.getFlightLog();
REQUIRE(flightLog.size() > 0);
double maxAltitude = 0.0;
double landingAltitude = flightLog.back().getPosition()[2];
for (const auto& state : flightLog) {
double alt = state.getPosition()[2];
if (alt > maxAltitude) {
maxAltitude = alt;
}
}
REQUIRE(maxAltitude > 0.0); // Rocket climbed
REQUIRE(landingAltitude <= 1.0); // Rocket landed (near ground)
// Optional: check that apogee was reached early in flight
size_t apogeeIndex = 0;
for (size_t i = 1; i < flightLog.size(); ++i) {
if (flightLog[i].getPosition()[2] < flightLog[i-1].getPosition()[2]) {
apogeeIndex = i - 1;
break;
}
}
REQUIRE(apogeeIndex > 0); // Apogee was found
}
}

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// tests/test_stage.cpp
#include "vendor/catch_amalgamated.hpp"
#include "Stage.h"
#include "Motor.h"
#include "Airframe.h"
#include "FinSet.h"
#include "RecoverySystem.h"
#include "Environment.h"
// Bring Approx into scope
using Catch::Approx;
TEST_CASE("Stage basic behavior", "[stage]") {
Stage stage("TestStage");
SECTION("Stage initially has zero mass and no components") {
REQUIRE(stage.getTotalMass() == Approx(0.0));
REQUIRE(stage.getTotalPropellantMass() == Approx(0.0));
REQUIRE(stage.getAirframeLength() == Approx(0.0));
REQUIRE(stage.calculateNormalForceCoefficient() == Approx(0.0));
}
SECTION("Adding a motor updates propellant mass") {
auto motor = std::make_unique<Motor>("TestMotor", 0.05, 50.0); // 50 grams
stage.addMotor(std::move(motor));
stage.updateMassProperties();
REQUIRE(stage.getTotalPropellantMass() > 0.0);
REQUIRE(stage.getTotalMass() > 0.0);
}
SECTION("Adding an airframe updates dry mass and length") {
auto airframe = std::make_unique<Airframe>("BodyTube",
1.2, // length meters
0.1, // diameter meters
1.5, // dry mass kg
0.75 // drag coefficient
);
double expectedLength = airframe->getLength();
double expectedMass = airframe->getDryMass();
stage.setAirframe(std::move(airframe));
stage.updateMassProperties();
REQUIRE(stage.getAirframeLength() == Approx(expectedLength));
REQUIRE(stage.getTotalMass() >= Approx(expectedMass));
}
SECTION("Adding fins provides stability contribution") {
auto fins = std::make_unique<FinSet>("BasicFins",
3, // 3 fins
0.15, // root chord
0.10, // tip chord
0.10, // span
0.05, // sweep length
0.003, // thickness
500.0 // material density
);
stage.setFinSet(std::move(fins));
REQUIRE(stage.calculateNormalForceCoefficient() > 0.0);
}
SECTION("Stage thrust calculation sums motor thrusts") {
auto motor1 = std::make_unique<Motor>("Motor1", 0.03, 30.0);
auto motor2 = std::make_unique<Motor>("Motor2", 0.02, 20.0);
motor1->addThrustDataPoint(0.0, 0.0);
motor1->addThrustDataPoint(0.5, 50.0);
motor1->addThrustDataPoint(1.0, 0.0);
motor2->addThrustDataPoint(0.0, 0.0);
motor2->addThrustDataPoint(0.5, 30.0);
motor2->addThrustDataPoint(1.0, 0.0);
stage.addMotor(std::move(motor1));
stage.addMotor(std::move(motor2));
stage.updateMassProperties();
stage.prepareForFlight(*(new Environment(101325.0, 288.15, 0.0))); // Dummy environment
// Simulate a half second into flight
for (auto& motor : stage.getMotorsForTesting()) {
motor->update(0.5);
}
auto thrustVector = stage.getTotalThrust();
REQUIRE(thrustVector[2] > 0.0);
}
}

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