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October 2025�Sofia, Bulgaria

DESIGN AND EXPERIMENTAL INVESTIGATION OF A SOLID PROPELLANT ROCKET MOTOR

Ana-Maria Bogdanova

NaFSKI 2025

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INTRODUCTION

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THE LEGEND OF WAN HU

Illustration courtesy of United States Civil Air Patrol - NASA

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“IGNIS”: It’s Guaranteed to Not Ignite Suddenly

Michael

Tsanev

Alexandar

Kalaydjiev

Valentin

Shopov

Ralitsa

Velikova

Ana-Maria

Bogdanova

Snejka

Grigorova

Nikola

Kazlachev

Ivan

Tanev

Nikola

Dekovski

Alexander

Zdravkov

Aleksandar

Todorov

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“IGNIS” ROCKET

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ROCKET COMPONENTS

Payload

Parachute

SOLID ROCKET MOTOR

Recovery system electronics

Main controller and communication module

Fins

Nose cone

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NOSE CONE

Ellipsoid shape:

Semi-major axis: 150mm
Semi-minor axis: 37.5mm

Camera opening

Body insert

Material : ABS, 50% infill

Average Cd = 0.05 (subsonic)

Camera housing

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Pressure sensors - BME 280

Controller - Raspberry Pi Zero 2W

Camera Module 3 Wide

Li-Po 3.7 V 500mAh

5V Step-Up/Step-Down S13V15F5

PAYLOAD

Accelerometer - Grove IMU 9-DOF

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Pressure sensor

Microcontroller

Power port

Li battery

Pyro charge

💥

Fired when desired height is reached.

RECOVERY SYSTEM

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RECOVERY SYSTEM MECHANICS

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MAIN CONTROLLER

Battery

LoRa Radio 433MHz

GNSS Module ZOE-M8

9-DOF IMU

Microcontroller ESP32S3

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GRAPHICAL INTERFACE AND COMMUNICATION

LoRa Radio 433MHz

Arduino Uno

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Number of fins: 3 (azimuthal symmetry)

Fins parameters: �Semi-span – most critical for rocket stability

Fins position:�15mm from the aft side of the rocket

Root chord

Tip chord

Semi-span

FINS

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LAUNCH PAD

3-point truss

Base

Launch rail

IGNIS Rocket

Truss height: 2000 mm

Rocket height: 1125 mm

Rail height: 3000 mm

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ROCKET MOTOR

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THEORETICAL PRINCIPLES

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MAIN COMPONENTS OF A SOLID ROCKET MOTOR

Nozzle

Combustion chamber

Insulation

Igniter system

Propellant

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PARAMETERS OF THE MOTOR

Thrust:

Specific impulse:

Total impulse:

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BASIC PRINCIPLES OF NOZZLE OPERATION

RESISTOJET

ARCJET

ION

Relationship between velocity change and cross-sectional area variation:

The function of the nozzle can be understood by examining the relationship between the change in cross-sectional area and the change in flow velocity as a function of the Mach number. For flow velocities below the speed of sound (M < 1), the relative changes in velocity 𝑑𝑣/𝑣and cross-sectional area 𝑑𝐴/𝐴have opposite signs. This means that in order to accelerate the flow (dv > 0), we must decrease the cross-sectional area (dA < 0) through which the fluid moves. On the other hand, at supersonic speeds (M > 1), the changes in velocity and area occur in the same direction — to increase the flow velocity (dv > 0), the cross-sectional area must be increased (dA > 0). In an efficient nozzle, the subsonic gas flow from the combustion chamber is accelerated in the converging section, reaches M = 1 at the throat, and continues to accelerate to supersonic speeds in the diverging section.

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TYPES OF EXHAUST FLOW

RESISTOJET

ARCJET

ION

Shock wave inside the nozzle

Overexpanded

flow

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RESISTOJET

ARCJET

ION

Underexpanded flow

Preferred flow mode

TYPES OF EXHAUST FLOW

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PROPELLANT AND BURN RATE MODEL

Potassium nitrate

(KNO₃)

Sorbitol

(C₆H₁₄O₆)

KNSB

65%

35 %

A_b – burning area

r – burn rate

a – burn rate coefficient

n – pressure exponent

P_ch – chamber pressure

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MODEL APPROXIMATIONS

The working fluid is homogeneous and compressible.

The working fluid obeys the ideal gas law.

All processes are adiabatic.

Wall friction is negligible.

Chemical equilibrium is established within the combustion chamber, and no chemical reactions occur in the nozzle.

The flow is one-dimensional.

The expansion of the flow in the nozzle is steady, without shockwaves and significant changes of gas properties.

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THERMODYNAMIC CHARACTERISTICS OF THE PROPELLANT

Adiabatic index

Mol numbers of gasses and condensed-phase particles

Reaction products and mol numbers

Adiabatic flame temperature

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THEORETICAL CHAMBER PRESSURE

After propellant depletion:

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THEORETICAL THRUST PROFILE

Thrust

Chamber pressure

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EXPERIMENT

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ROCKET MOTOR DESIGN

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ROCKET NOZZLE DESIGN

Throat cross section:

Exit cross section :

Convergence angle: 60˚

Divergence angle : 20 ˚

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PROPELLANT

D = 45 mm

d = 14 mm

L = 70 mm

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IGNITER SYSTEM

Battery connector

Switch

Battery (9V)

Conducting wire

Nichrome wire

Paper tube, covered with rocket fuel

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EXPERIMENTAL SETUP

Test stand

Load cell

ARDUINO MEGA

Analog-to-Digital Converter (ADC) for Weigh Scales HX711

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EXPERIMENT

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THRUST COMPARISON: THEORY VS EXPERIMENT

Experiment

Theory

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COMPARISON BETWEEN THEORY AND EXPERIMENT

PARAMETER	COMPUTATION METHOD	THEORY	EXP.	EXP./THEORY
	Maximum value for each graph	275	245	89.1 %
		522	439	84.0 %
	The duration in each graph when thrust is > 0	2.11	3.5	165.9 %
		247	125	50.7 %
		150	127	84.8 %
		0.168	0.100	59.8 %
		1475	1250	84.7 %

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TRAJECTORY AND FLIGHT

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EXPECTED TRAJECTORY

Apogee: 1010 m

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NEXT STEPS

Design validation and testing
Flight preparation
Flight
Data post-processing

Obtaining video and using software stabilization
Obtaining accelerometer and gyroscope data and estimating attitude and trajectory

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THANK YOU FOR YOUR ATTENTION!

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Content

Introduction

“IGNIS” Rocket

Rocket motor

Theoretical principles
Propellant and burn rate model
Model approximations
Thermodynamic characteristics of the propellant
Theoretical chamber pressure and thrust profile
Rocket motor design
Experiment