Rocket-Graph-Maker

1D Model Rocket Flight Simulator

A lightweight Python simulation of a model rocket’s vertical (1D) flight from launch to apogee. It uses a tabulated motor thrust curve, accounts for propellant mass loss and quadratic aerodynamic drag, and integrates the equations of motion with a 4th-order Runge-Kutta (RK4) solver.

Features

• Tabulated thrust curve with linear interpolation
• Time-varying mass, based on the fraction of total impulse delivered
• Quadratic drag that always opposes velocity
• RK4 integration of height and velocity
• Reports burnout time, max velocity, apogee height, time to apogee, and total impulse
• Saves a 4-panel plot of height, velocity, acceleration, and mass vs. time

Requirements

• Python 3.9+
• NumPy 2.0 or newer (the script uses np.trapezoid, which does not exist in older versions; on NumPy 1.x, replace it with np.trapz)
• Matplotlib

pip install numpy matplotlib

Usage

python rocket_sim.py

(Replace rocket_sim.py with whatever you named the file.)

The script prints a summary to the console and writes rocket_flight.png to the working directory.

Burnout time:   ...
Max velocity:   ...
Apogee height:  ...
Time to apogee: ...
Total impulse:  ...

Configuring Your Rocket

Edit the constants near the top of the script.

Parameter	Default	Units	Description
MASS_EMPTY	0.030	kg	Rocket structure plus motor casing, no propellant
MASS_PROPELLANT	0.015	kg	Propellant burned during flight
DRAG_COEFFICIENT	0.75	–	Drag coefficient (Cd)
FRONTAL_AREA	0.0015	m²	Cross-sectional area of the rocket
AIR_DENSITY	1.225	kg/m³	Air density at sea level
G	9.81	m/s²	Gravitational acceleration
DT	0.01	s	Integration time step

Thrust curve

THRUST_CURVE is an array of [time (s), thrust (N)] pairs. Replace it with data for your motor (manufacturer data sheets and the ThrustCurve.org database list these). The first point should be at t = 0 and the last point should have zero thrust, since the last time value is treated as burn time.

How It Works

State: [height, velocity]

Forces:

F_net = F_thrust(t) - F_drag(v) - m(t) * g
F_drag = 0.5 * rho * v * |v| * Cd * A
a = F_net / m(t)

Mass model: Propellant is assumed to burn in proportion to the impulse delivered so far:

m(t) = m_empty + m_propellant * (1 - cumulative_impulse(t) / total_impulse)

Integration: Each step of size DT uses classic RK4 on d/dt [h, v] = [v, a(t, v)].

Stopping condition: The simulation ends at apogee, defined as the first time after burnout when velocity drops to zero or below. There is also a 60 s safety cutoff.

Output

rocket_flight.png contains four stacked plots sharing a time axis:

1. Height (m), with apogee in the title
2. Velocity (m/s)
3. Acceleration (m/s²)
4. Mass (kg)

A dashed green line marks motor burnout on each plot.

Assumptions and Limitations

• 1D only. Vertical flight with no wind, weathercocking, or angle of attack.
• Constant Cd and air density. No variation with Mach number or altitude, which is reasonable for low-altitude model rockets.
• Constant frontal area and no parachute. The simulation stops at apogee and does not model descent or recovery.
• No launch rail or lug friction.
• Linear impulse-based propellant burn. Real motors may deviate.
• Results are only as good as your inputs. Validate against a tool such as OpenRocket or real flight data before relying on them.

Possible Improvements

• Descent and parachute deployment phases
• Altitude-dependent air density (ISA atmosphere model)
• Mach-dependent drag coefficient
• Load thrust curves directly from .eng files
• Command-line arguments for rocket parameters
• Comparison against OpenRocket or flight-computer data

License

License

This project is licensed under the MIT License.