People’s Democratic Republic of Algeria
Ministry of Higher Education and Scientific Research
Ecole Nationale Polytechnique de Constantine
Project Report
SkyBridge Rocket
Rocketry Competition Report – 5th Edition
CIAAR, Blida
Students:
Bendjeddou Ibrahim Salah Eddine
BenHamlaoui Abderrahmane
Bouharati Badr Elmounir
Chebel Ouissal
Chouali Ouissal Malak
Hamlaoui Ikram
Kehal Isra
Supervisors:
Dr. Kharoua Nabil
Dr. Mouadji Youcef
Pr. Khezzar Lyes
Dr. Teniou Samir
Academic Year: 2024/2025
ENPC Rocket Project Report © 2024 is licensed under CC BY-NC 4.0
“Everyone knew it was impossible,
until a fool who didn’t know
came along and did it.”
— Albert Einstein
Contents
List of Figures
5
List of Tables
8
Notation
9
Acknowledgments
10
Abstract
11
1
Competition and Team Overview
1.1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
1.2 The CIAA Rocketry . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
1.3 Aims & Objectives . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
1.4 Our Team . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
13
13
14
14
15
2
Structural system: Rocket Design and Considerations
2.1 Our Rocket Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
2.1.1 10K feet Rocket . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
2.1.2 The Low Altitude Rocket . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
2.2 Rocket Design Considerations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
2.2.1 Sizing Calculations for Rocket Components . . . . . . . . . . . . . . . . . . . . . . . . . . .
16
16
16
17
18
18
3
Payload
3.1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.2 System Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.2.1 System’s description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.2.2 Payload specifications and mission . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.2.3 Payload block diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.3 Hardware Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.3.1 Data processing unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.3.2 Sensors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.3.3 GPS module . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.3.4 SD Card Module . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.4 System Integration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.4.1 Schematics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.4.2 PCB design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.5 Tests and validation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.5.1 Sensor calibration and tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.5.2 PCB circuit continuity test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.5.3 PCB stability test against vibrations and chock . . . . . . . . . . . . . . . . . . . . . . . . . .
3.6 Conclusion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
20
20
20
20
20
21
21
21
22
25
25
26
26
27
29
29
30
30
31
2
CONTENTS
3
4
Recovery & Avionics
4.1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.2 System Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.2.1 System’s description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.2.2 Avionics block diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.3 Hardware components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.3.1 Flight computer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.3.2 Ground station . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.3.3 Sensors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.3.4 Deployment trigger circuit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.4 System Integration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.4.1 Schematics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.4.2 PCB design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.4.3 EMI considerations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.4.4 Isolation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.5 Software . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.5.1 Flight states/phases . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.5.2 Altitude calculation method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.5.3 Deployment algorithm . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.6 Testing and Validation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.6.1 Sensor calibration and tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.6.2 Deployment circuit test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.7 Conslusion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
32
32
32
32
33
34
34
35
35
35
36
36
38
38
39
39
39
40
40
40
40
42
43
5
Structural Strength Analysis
5.1 Main Rocket . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5.1.1 Fin Flutter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5.1.2 Aerodynamic Loading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5.1.3 Flight loads . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5.1.4 Calculating parachute size and resistance . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5.2 Structural Strength Calculations Low Altitude Rocket . . . . . . . . . . . . . . . . . . . . . . . . . .
44
44
44
45
46
52
53
6
Softwares: Analysis and Simulations
6.1 Openrocket . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.1.1 Thrust curve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.2 Rocksim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.2.1 Thrust curve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.3 Openrocket Vs Rocksim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.3.1 Similarities . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.3.2 Differencies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.4 SOLIDWORKS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.4.1 Simulation of the nosecone . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.4.2 Loads and Fixtures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.4.3 Study Result . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.4.4 Simulation of the fins . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.4.5 Loads and Fixtures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.4.6 Study Result . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.5 Computational fluid dynamics (CFD) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.5.1 Grid Generation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.5.2 Mesh Effect . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.5.3 Solution settings and boundary conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.5.4 Boundary conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.5.5 Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.5.6 Conclusion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
55
55
56
56
57
57
57
58
59
59
60
60
64
64
65
67
67
70
71
71
71
74
CONTENTS
4
7
Manufacturing
7.1 Manufacturing the Main Rocket . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.1.1 Manufacturing of the Body tube . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.1.2 Nose cone . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.1.3 Fins . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.1.4 Centering rings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.1.5 Bulkheads . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.1.6 Coupler . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.1.7 Parachute . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.2 Manufacturing the Low altitude Rocket . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.2.1 Manufacturing of the Body tube . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.2.2 Fins . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.2.3 Centering rings and bulkheads . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.2.4 The Nose Cone . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.2.5 Parachute . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.3 Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.3.1 Body Tube Coupling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.3.2 Parachute Packing and Placement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.3.3 Shock Cord Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.3.4 Electronics Integration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.3.5 Black Powder Ejection System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.3.6 Final Assembly Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
75
75
75
78
78
81
81
82
82
83
83
83
85
86
86
87
87
87
88
88
88
89
8
Tests and Validations
8.1 Fiberglass Thermal Resistance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
8.2 Ejection charge test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
8.3 Deployment test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
8.4 Tensile tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
8.4.1 PLA & PETG . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
8.4.2 Fiberglass composite . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
8.4.3 Shock cord and shroud lines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
92
92
92
93
93
94
95
96
Conclusion
Project Self-Evaluation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Conclusion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
97
97
97
Bibliography
98
APPENDIX A: Drawing of the design ( Low altitude rocket )
99
APPENDIX B: Virtual maquette of the main rocket
104
APPENDIX C: Barrowman method for calculating center of pressure
107
APPENDIX D: Static Stability Criteria
108
APPENDIX E: Rocket Materials List
109
List of Figures
1.1
Team members . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
15
2.1
2.2
2.3
2.4
10k ft rocket design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Low altitude rocket . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Fin dimensions (Rick Hanke 2010) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Parachute area configurations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
17
18
19
19
3.1 Payload Block Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.2 Arduino Mega 2560 micro-controller (Sensor Embedded 2024) . . . . . . . . . . . . . . . . . . . . .
3.3 AHT20 humidity sensor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.4 BMP280 Pressure sensor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.5 BMP280 and AHT20 sensor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.6 MPU6050 Accelerometer and Gyroscope . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.7 The MQ135 sensor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.8 GPS module . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.9 SD Card Module . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.10 Payload schematics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.11 Payload prototype circuit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.12 PCB . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.13 3D Front . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.14 3D Back . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.15 Payload PCB first view . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.16 Payload PCB second view . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.17 Payload circuit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
21
21
22
23
23
24
24
25
25
26
27
27
28
28
28
28
29
4.1 Dual deployment stages . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.2 Avionics block diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.3 Heltec LoRa32 V2 micro-controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.4 The IRFZ44N MOSFET . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.5 Optocoupler . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.6 E-matches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.7 Deployment system Schematics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.8 Deployment system circuit: transmitter and receiver . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.9 PCB Design for the deployment system circuit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.10 3D representation of the PCB for the deployment system circuit . . . . . . . . . . . . . . . . . . . .
4.11 PCB of the deployment system circuit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.12 The relationship between altitude and pressure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.13 Deployment algorithm . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.14 Measured altitude from sea level . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.15 Filtered measured altitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.16 Apogee detection test results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
33
33
34
35
36
36
37
37
38
38
39
40
41
42
42
43
5.1
5.2
5.3
5.4
5.5
46
46
49
50
51
REFPROP properties at maximum velocity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Angle of Attack . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Aero and inertia loads acting on rocket body in flight . . . . . . . . . . . . . . . . . . . . . . . . . .
Effects of Shear and Bending Moment on rocket . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Shear force Vs Distance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5
LIST OF FIGURES
5.6
6
Bending moment Vs Distance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
51
6.1 Skybridge design in Openrocket . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.2 Skybridge design in Openrocket . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.3 Thrust curve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.4 Final design Rocksim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.5 Simulations on Rocksim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.6 Thrust curve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.7 Altitude Vs Time or (Trajectory of flight) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.8 Units used in static simulation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.9 model nosecone . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.10 Different properties of the nosecone . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.11 Loads and Fixtures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.12 nosecone stress . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.13 nosecone stress . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.14 nosecone strain . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.15 Safety factor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.16 the fins . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.17 Material proprieties . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.18 Loads and Fixtures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.19 Fines stress . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.20 Fines stress . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.21 Fins strain . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.22 Fins factor of safety . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.23 Mesh 1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.24 Mesh 2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.25 Mesh 3 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.26 Drag coefficient from OpenRocket . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.27 Drag force graph . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.28 Drag Coefficient . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.29 Drag force from Ansys Fluent (Mesh 3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.30 Velocity contour . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.31 Velocity at the tip of the nose cone . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.32 Static temperature contour . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.33 Static pressure contour . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
55
56
56
56
57
57
58
59
59
60
60
61
62
63
63
64
64
65
65
66
66
67
68
68
69
70
70
71
71
72
72
73
74
7.1 The PVC pipe cutting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.2 The half of the body tube . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.3 Release agent . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.4 Gel Coat application . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.5 Resin preparation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.6 Resin application . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.7 Acrylic mastic sealant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.8 Nose cone part . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.9 Nose cone assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.10 The piece of Fiberglass . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.11 Cutting the fins . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.12 Final result of the fins . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.13 Tang slotted airframe . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.14 Fins attachment using shear pin . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.15 Fins attachment using Fiberglass . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.16 The mold of centering rings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.17 Final result of the centering rings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.18 Hole cutter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.19 Cutting the parachute parts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.20 Parachute Sewing Process . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.21 Final Outcome of the Parachute Sewing Process . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.22 Leveling the body tube using a Level Bubble . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
76
76
76
77
77
77
77
78
78
79
79
79
79
80
80
81
81
82
83
83
83
84
LIST OF FIGURES
7
7.23 Fins cutting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.24 Tang slotted airframe . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.25 Glowing the fins . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.26 Centering rings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.27 Bulkhead . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.28 Glowing the fins . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.29 Nose Cone 3D printed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.30 The coupler . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.31 Shock cord installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.32 Glue sticks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.33 Assembled rocket . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.34 Painting process . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
84
84
85
85
85
86
86
87
88
89
90
91
8.1
8.2
8.3
8.4
8.5
8.6
8.7
8.8
92
93
94
94
95
95
95
96
Fiberglass Thermal Resistance Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Ejection charge test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Tensile testing machine . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
PLA & PETG test specimens . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Force variation over time for PLA . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Force variation over time for PETG . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Fiberglass composite test specimens . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Shroud line test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
List of Tables
3.1
3.2
3.3
3.4
3.5
Specifications of the ATmega2560 Microcontroller . . . . . . . . . . . . . . . . . . . . . . . . . . . .
AHT20 sensor Specifications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Specifications of the BMP280 sensor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Specifications of the Sensor MPU6050 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Sensor Testing and Calibration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
22
22
23
24
30
4.1
Specifications of the Device . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
35
5.1
Mechanical properties of the composite material . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
45
6.1
Drag force and Cd for different meshes. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
70
8.1
Rocket Materials List: Mechanical and Electronic Components . . . . . . . . . . . . . . . . . . . . . 109
8
Notation
Notation
Signification
Unit
A
a
S
Ar
AR
CD
CL
CNα
Cr
Ct
Cx
D
E
Fd
g
G
h
M
Nx
P
P0
q
ρ
Sf
T
U
V
Vd
Vo
Vf
Vmax
α
λ
Frontal area of the rocket
Speed of sound at altitude of maximum velocity
Surface area of the parachute
Reference area
Aspect ratio
Drag coefficient
Lift coefficient
Normal force coefficient
Root chord length
Tip chord length
Drag coefficient of the parachute
Diameter
Young modulus
Drag force
Gravity
Shear modulus
Altitude where maximum velocity was reached
Total mass of the rocket
Normal force
Air pressure at the altitude of maximum velocity
Air pressure at sea level
Dynamic pressure
Air density
Fin area
Thrust
Gust velocity
Vehicle’s velocity
Descent velocity of the rocket
Opening velocity
Flutter speed
Maximum velocity
Angle of attack
Fin taper ratio
m2
m/s
m2
m2
/
/
/
/
m
m
/
m
GPa
N
m/s2
Pa
m
kg
N
Pa
Pa
N/m2
kg/m3
m2
N
m/s
m/s
m/s
m/s
m/s
m/s
rad
/
9
Acknowledgment
Before delving into the details of our project, we would like to express
our gratitude to Allah for granting us the strength and perseverance to
bring this project to fruition. Without his grace, we would not have
come this far.
We are deeply grateful to all those who have contributed, directly or
indirectly, to the success of this endeavor.
We extend our heartfelt appreciation to the National Polytechnic School
of Constantine for providing an excellent educational foundation and the
necessary resources to pursue our passion for rocketry.
Our profound appreciation goes to our supervisors, Mr. Kharoua Mr.
Mouadji, Mr. Khezzar, and Mr. Teniou, for their invaluable guidance,
unwavering support, and dedicated efforts.
We would also like to acknowledge our friends—Cherifi Yahya Mourad, Gasmi
Ahmed Anis, and Radi Wafa—who offered their assistance during various
stages of this project. Their support has made this project a truly
collaborative effort.
Additionally, we are grateful to the staff members of the National
Polytechnic School of Constantine who facilitated our work.
A special mention goes to Mr. Kherrat, who introduced us to the
fascinating world of rocketry, igniting our passion for this field.
Finally, we extend our thanks to the esteemed members of the jury for
honoring us by reviewing and evaluating our work.
10
Abstract
This comprehensive report outlines the design, simulation, and structural analysis of a high-performance rocket tailored specifically for the CIAA Rocketry competition.
Leveraging a cutting-edge software suite, our methodology incorporates SOLIDWORKS for initial 3D modeling,
OpenRocket for preliminary aerodynamic assessments, and ANSYS Fluent for computational fluid dynamics (CFD)
simulations.
The design process is guided by the dual objectives of adhering to CIAA Rocketry regulations and optimizing both
aerodynamic efficiency and structural resilience. OpenRocket is utilized to explore a range of fin configurations and
the identification of optimal geometries. ANSYS Fluent simulations provide a detailed analysis of the rocket’s behavior under launch loads, with visualizations of flow properties in order to facilitate the identification of potential
weak points and informing design refinements.
The final design, incorporating these insights, will be presented in detail, including precise dimensions, material
selection, and a comprehensive Bill of Materials, thereby providing a complete roadmap for the construction of a
competitive rocket for CIAA Rocketry.
Keywords: High-performance rocket, structural analysis, CIAA Rocketry competition, SOLIDWORKS, OpenRocket,
ANSYS Fluent, CFD, Propulsion, Flight phases.
Résumé
Ce rapport complet décrit la conception, la simulation et l’analyse structurelle d’une fusée haute performance conçue
spécifiquement pour CIAA Rocketry.
S’appuyant sur une suite logicielle de pointe, notre méthodologie intègre SOLIDWORKS pour la modélisation 3D initiale, OpenRocket pour les évaluations aérodynamiques préliminaires et ANSYS Fluent pour les simulations avancées
de dynamique des fluides numérique (CFD).
Le processus de conception est guidé par le double objectif de respecter les réglementations de la CIAA Rocketry
et d’optimiser à la fois l’efficacité aérodynamique et la résilience structurelle. OpenRocket est utilisé pour explorer
une gamme de configurations d’ailerons et l’identification des géométries optimales. Les simulations ANSYS Fluent
fournissent une analyse détaillée du comportement de la fusée sous les charges de lancement, avec des visualisations
des propriétés de l’écoulement afin de faciliter l’identification des points faibles potentiels et d’informer sur les raffinements de la conception.
La conception finale, incorporant ces idées, sera présentée en détail, y compris les dimensions précises, la sélection
des matériaux, et une nomenclature complète, fournissant ainsi une feuille de route complète pour la construction
d’une fusée compétitive pour le CIAA Rocketry.
Mots-clés : Fusée haute performance, analyse structurelle, compétition CIAA Rocketry, SOLIDWORKS, OpenRocket,
ANSYS Fluent, CFD, propulsion, phases de vol.
11
12
LIST OF TABLES
‘jÊÓ
QK Q®JË@ @ Yë Qk. ñK
CIAA é®K. A‚ÖÏ A’J’k ÕÔ’Ó Z@X B@ úÍA« pðPA’Ë úξJêË@ ÉJÊjJË@ð èA¿AjÖÏ @ð ÕæҒJË@ ÉÓA‚Ë@
. Rocketry
YÒJÊË SOLIDWORKS l× AKQK Ð@YjJƒ@ AJJ jîDÓ áҒJK , èPñ¢JÓ HAJ
× áÓ èXA®JƒBAK
m.× QK. é«ñÒm
éJ KCK éJ Ëð B@ ék
.
.
.
.
. .
×
×
ÉK@ñ‚Ë@ A¾JÓAJK X èA¿AjÖÏ ANSYS Fluent l . AKQK. ð , éJËð B@ éJºJÓAJK XðQK B@ HAÒJJ®JÊË OpenRocket l . AKQK. ð ,XAªK. B@
® JÖÏ @ éJ K A‚mÌ '@
. (CFD) éÓY
.
ƒ éÖ ß A¯ð
X@ñÖÏ @ PAJJk@ð é®J ¯YË@
XAªK B@ ½Ë Y» áҒJK ø YË@ ,ùKAîDË@ ÕæҒJË@ Õç' Y® K ÕæJ ƒ
Q¯ñK úÍAJËAK. ð ,X@ñÒÊË éÊÓA
.
CIAAR é®K. A‚ÖÏ ú愯AJK pðPA’Ë éÊÓA¿ ‡K Q£ é£PAg
ANSYS , SOLIDWORKS , OpenRocket , úξJêË@ ÉJÊjJË@ ,©¯YË@ ,Z@X B@ úÍA« pðPA“ : éJkAJ®ÖÏ @ HAÒʾË@.
. CIAAR é®K . A‚Ó , CFD , Fluent
Chapter 1
Competition and Team Overview
1.1
Introduction
“It is not Rocket Science” is an idiom frequently used to convey that something is not overly difficult to understand
or complicated to accomplish. It refers to a relatively simple concept that doesn’t require specialized knowledge or
expertise. What is rocket science? Why is it considered hard? And why do we need rockets?
Rocket science, also known as astronautics or simply rocketry, is a field of engineering focused on the design, operation, maintenance, and development of vehicles capable of flight in outer space. It covers a broad range of disciplines,
from advanced mathematics, chemistry, material science, aerodynamics, physics, logistics, systems engineering, and
yes, even politics.
While all fields of engineering are considered challenging, rocketry has earned particular attention for various factors,
one being a balancing act, that involves precise calculations, technical skills, trajectory planning, fuel management,
and selection of high-endurance materials. In this high-stakes environment, even a single misstep can lead to catastrophic consequences—both human and economic. The line between a successful launch and a disastrous failure is
often razor-thin.
Given these challenges, one might wonder, is it worth it? If it were not for rockets, we wouldn’t be able to use our cell
phones, watch a lot of our favorite television shows, find out the weather forecast, navigate with Global Positioning
System (GPS), or explore our solar system—just to name a few.
13
CHAPTER 1. COMPETITION AND TEAM OVERVIEW
1.2
14
The CIAA Rocketry
It is clear that building a rocket requires specialized knowledge and often the collaboration of hundreds of people.
However, there is an accessible way for enthusiasts to engage with this fascinating field: model rocketry.
This practice allows individuals to safely build and launch rockets on a smaller scale, bridging the gap between amateur hobbyists, professional scientists, and engineers.
This brings us to the CIAA Rocketry competition, which has provided us and many Algerian students over the years
with the chance to design and build our own model rockets, and the reason we are writing this report today.
After a hiatus due to the COVID-19 pandemic, the competition’s return presents a prime opportunity for us to represent our school in the 2024 edition.
1.3
Aims & Objectives
As Algerian students, we frequently encounter the phrase "Are you launching rockets?" used to undermine our academic efforts. This project aims to challenge that perspective by demonstrating that dedicated study can lead to
impressive achievements, such as building rockets.
Our primary objective is to prepare for the CIAA Rocketry’s 5th edition by ensuring we meet all technical requirements while designing and constructing a model rocket capable of successful flight.
Additionally, this project highlights the importance of teamwork, as we work collaboratively with team members,
supervisors, and external experts to ensure the successful completion of the project.
Finally, through this endeavor, we aimed to revive interest in amateur rocketry within our school community, by
carefully documenting our findings through presentations and written materials, contributing to the wider body of
knowledge in amateur rocketry.
CHAPTER 1. COMPETITION AND TEAM OVERVIEW
1.4
15
Our Team
The "Skybridge" team was established in December 2023 to represent l’École Nationale Polytechnique de Constantine in the fifth edition of the CIAA Rocketry competition. This initiative aims to rekindle the school’s enthusiasm
for amateur rocketry, having previously participated only in the 2nd edition of the event. The team comprises seven
master’s students along with four faculty advisors from both the Mechanical Engineering Department and the Electrotechnical and Automation Department.
Figure 1.1: Team members
Benhamlaoui Abderrahmene, Bouharati Badr Elmounir, and Kehal Isra, from the EEA department, have been in total
charge of the electronics system and the payload under the supervision of Dr. Teniou Samir.
Bendjeddou Ibrahim, Chebel Ouissal, Chebel Ouissal Malak, and Hamlaoui Ikram are the ones in charge of the structural system, mainly its manufacturing, structural analysis under the supervision of Dr Youcef Mouadji, and all kinds
of simulations under the supervision of Dr Kharoua and Pr Khezzar.
Chapter 2
Structural system: Rocket Design and
Considerations
A rocket consists essentially of four systems:
• Structural system
• Propulsion system
• Payload
• Recovery system
Additionally, rockets may incorporate a fifth system for guidance, typically comprising radars and computers to
ensure precise navigation. (Rockets 101 | National Geographic, 2019)
In this report, we will focus on three of these crucial systems: structural, payload, and recovery.
The propulsion system includes all parts of the rocket engine, such as propellants, power head, and rocket nozzle.
There are two main types of rocket engines: liquid engines, which store fuel and oxidizer separately, and solid engines,
where fuel and oxidizer are combined and packed into a solid cylinder. (Propulsion System, 2023)
In Model rocketry, which is our focus, solid motors are commonly used. Since we will be using a commercially
purchased motor, there is no need to explore the details of the propulsion system further for this discussion.
2.1
Our Rocket Design
2.1.1
10K feet Rocket
The main rocket, or the 10K feet rocket, consists of three main sections: nose cone, forward airframe, and aft airframe.
It has a 19.4 cm outer diameter and is 245 cm long. It weighs 28.59 kg at liftoff, and is expected to reach an altitude
of 9878.61 ft AGL (3011 m). The rocket’s motor is N-class.
The minimum static stability is 2.18 Cal as determined by Openrocket.
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CHAPTER 2. STRUCTURAL SYSTEM: ROCKET DESIGN AND CONSIDERATIONS
17
Figure 2.1: 10k ft rocket design
The nose cone section houses the payload.
The forward airframe houses the drogue parachute, its shock cords, and the mass component beneath the parachute
The mass component beneath the parachute simulates a piston that is used to eject the nose cone while also protecting the parachute.
The aft airframe contains the main parachute and its shock cords. Similar to the forward airframe, the mass component here also simulates a piston, which is used to both eject and protect the main parachute. Additionally, the aft
airframe houses the motor, motor mount, and fins.
The forward and aft airframes are connected by a coupler, which serves as a housing for the electronic components,
also known as the recovery system.
2.1.2
The Low Altitude Rocket
The small rocket, or the low altitude rocket, consists 2 sections : nose cone, and an airframe.
It has an 11 cm outer diameter and is 123 cm long, it weighs 3.342 kg at liftoff, and is expected to reach an altitude of
1539 m. The rocket’s motor is J-class.
The minimum static stability is 1.34 Cal as determined by Openrocket.
CHAPTER 2. STRUCTURAL SYSTEM: ROCKET DESIGN AND CONSIDERATIONS
18
Figure 2.2: Low altitude rocket
This rocket doesn’t have a payload, so the nose cone section is empty, it also has a single airframe that houses the
parachute, the motor, the motor mount, and the fins.
2.2
Rocket Design Considerations
2.2.1
Sizing Calculations for Rocket Components
a. Diameter and Length of The Rocket
Rocket’s diameter depends on the largest motor we might want to fly in it. We can always adapt down, but can never
adapt never up.
The ratio of rocket length to diameter, should be from 10:1 to 20:1. For example, a 16 cm diameter rocket would
mean a length of 160 -320 cm. (Rules of Thumb.)
b. Fin Sizing
• Span (S) / Height: The distance from the leading edge to the trailing edge of the fin. The height should be
approximately 5 to 12.5% of the rocket body length, measured from the base. (Model Rocket Guide.)
• Root Chord (Cr ): The width of the fin where it attaches to the rocket body. Typically, the root chord is around
1-2 of the body diameter. (Gasmire.)
• Tip Chord (Ct ): The width of the fin at the tip. Typically, around 70-80% of the root chord or approximately
1 diameter length (Sizing Fins | STAR Public, 2020).
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CHAPTER 2. STRUCTURAL SYSTEM: ROCKET DESIGN AND CONSIDERATIONS
• Sweep Length: This is the distance from the leading edge of the root chord to the leading edge of the tip chord.
• Sweep Angle (θ): The angle between the leading edge and the rocket body. This can be calculated using
trigonometry:
Sweep Angle = tan
−1
Sweep Length
Height
• Thickness (T): The thickness of the fin, which depends on the material strength and aerodynamic requirements.
Figure 2.3: Fin dimensions (Rick Hanke 2010)
c. Parachute sizing
The surface area of a parachute required for effective deployment can be calculated using the following formula:
Sp =
2M g
ρCX Vd
We generally aim for a descent speed of 10 m/s.
Figure 2.4: Parachute area configurations
(2.1)
Chapter 3
Payload
3.1
Introduction
Payloads have been integral to the history of rockets, serving as a powerful testament to human ingenuity and innovation. Payloads have determined the purpose and impact of a rocket from the earliest incendiary devices, in ancient
rockets, to complex scientific instruments used in modern space missions.
The transition started from military usages, where rockets carried explosives, through the scientific phase in the 20th
century with atmospheric probes and later the first satellites (Sputnik 1). This change was the beginning of payloads
as simply a method for exploration, as well as communication and research. In subsequent years, rockets launched
interplanetary probes, manned spacecraft, and commercial satellites all carrying specialized payloads unique to mission goals.
Payloads today may include CubeSats, environmental sensors, biological experiments, or even space-tourism systems! This legacy inspires us in thinking through effective designs for our payloads that can achieve their functional
roles serving the missions of exploration and discovery.
In this chapter, we are going to explore in detail the payload used in our rocket.
3.2
System Overview
This section will give an overview of the payload, description , specification and mission.
3.2.1
System’s description
A rectangular payload of 10 cm long, 10 cm wide and 30 cm high weighing a total mass of 4 kg for the rocket was
used. It is housed inside the nose cone of the rocket and it will not be ejected due to the requirements of its mission.
Payload requirements for mission success dictate that the payload must be intact, undamaged and has never been
subjected to smoke resulted from the recovery system.
Though other details about the specific payload’s mission will be in the following subsection.
3.2.2
Payload specifications and mission
Our payload consists of a measurement unit which is a comprehensive data collection system designed for monitoring
of the aircraft’s performance and environmental conditions.
It incorporates an array of high-precision sensors including a BMP280 barometric pressure and temperature sensor,
used furthermore for altitude measurement; an AHT20 humidity sensor; an MQ135 gas sensor for air quality analysis
and pollution levels detection and a MPU6050 sensor for acceleration and angular velocity measurement. Data is
processed and stored in an SD card to be treated post flight, this operation is performed by the Arduino Mega 2560.
20
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CHAPTER 3. PAYLOAD
The system is powered by a LiPo battery and is programmed in the C++ programming language. Whereas, the data
gathered is visualized and examined using a different software.
3.2.3
Payload block diagram
In this subsection, we are providing a payload diagram for better understanding of its working principle and organization.
Figure 3.1: Payload Block Diagram
3.3
Hardware Components
This section provides an in-depth overview of the hardware integrated into the rocket payload, detailing the microcontroller, sensors, and electronic modules that work together to collect and store environmental data during the
flight. Where each component has been carefully selected to ensure reliable performance and accurate measurements,
contributing to the overall success of the mission.
3.3.1
Data processing unit
To process sensor readings, translate the data, and store flight information, a reliable microcontroller is essential. For
this purpose, we selected the Arduino Mega 2560.
The Arduino Mega 2560 is a powerful and versatile microcontroller board based on the ATmega2560. It is designed
for complex projects requiring a large number of I/O pins and higher processing capabilities compared to standard
Arduino boards.
Figure 3.2: Arduino Mega 2560 micro-controller (Sensor Embedded 2024)
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CHAPTER 3. PAYLOAD
• Technical specifications
Specification
Microcontroller
Operating Voltage
Input Voltage (recommended)
Digital I/O Pins
Analog Input Pins
DC Current per I/O Pin
DC Current for 3.3V Pin
Flash Memory
Clock Speed
Length
Width
Weight
Details
ATmega2560
5V
7-12V
54 (of which 15 provide PWM output)
16
20 mA
50 mA
256 KB (8 KB used by bootloader)
16 MHz
101.52 mm
53.3 mm
37 g
Table 3.1: Specifications of the ATmega2560 Microcontroller
3.3.2
Sensors
a. Humidity sensor
In order to measure the temperature and the humidity during the flight we used The AHT20, a high-performance
temperature and humidity sensor widely used in environmental monitoring applications. The AHT20 is an upgraded
version of the AHT10 sensor, offering improved accuracy and reliability with a Low power usage. It integrates a
capacitive humidity sensor and a high-precision thermistor into a single compact package.
Figure 3.3: AHT20 humidity sensor
• Sensor specifications
Parameter
Resolution
Accuracy
Measurement range
Input voltage
Current
Power consumption
Communication
Description
Humidity
Temperature
Humidity
Temperature
Humidity
Temperature
Min
Typical
Max
Typical
Average
Two-line digital interface, standard I2C protocol
Table 3.2: AHT20 sensor Specifications
Value
0.024
0.01
±2
±0.3
0 to 100
-40 to 85
2
3.3
5.5
23
3.3
Unit
%RH
°C
%RH
°C
%RH
°C
V
µA
µW
23
CHAPTER 3. PAYLOAD
b. Pressure sensor
For measuring pressure, temperature, and altitude, we selected the BMP280, a highly integrated environmental
sensor renowned for its precision in barometric pressure and temperature measurement, making it well-suited for
altitude determination in aerospace applications. As an upgraded version of the BMP180, the BMP280 delivers enhanced performance, reduced power consumption, and additional features, making it a popular choice for altimetry,
weather forecasting, and environmental monitoring.
Figure 3.4: BMP280 Pressure sensor
• Sensor specifications
Parameter
Resolution
Accuracy
Measurement range
Input voltage
Current
Power consumption
Communication
Description
Pressure
Temperature
Pressure
Temperature
Pressure (altitude)
Temperature
Min
Typical
Max
Typical
Average
Two-line digital interface, standard I2C protocol
Value
0.16
0.01
±1
±0.5
300 to 1100 (-500 to 9000)
-40 to 85
1.7
1.8
3.6
720
2.7
Unit
Pa
°C
Pa
°C
hPa (m)
°C
V
V
µA
µW
Table 3.3: Specifications of the BMP280 sensor
• The two previous sensors are implemented in one PCB :
Figure 3.5: BMP280 and AHT20 sensor
c. Accelerometer and Gyroscope
The MPU6050 is a popular 6-axis motion tracking sensor that combines a 3-axis gyroscope and a 3-axis accelerometer in a single chip. We used it to get precise measurements of the acceleration, velocity and orientation of
the rocket during the flight.
The MPU6050 is part of InvenSense’s family of motion sensors. Using an integrated Digital Motion Processor (DMP),
the motion data is directly processed on the chip, reducing the computational load on the host controller.
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CHAPTER 3. PAYLOAD
Figure 3.6: MPU6050 Accelerometer and Gyroscope
• Sensor specifications
Parameter
Full-Scale Range
Input voltage
Current
Power consumption
Communication
Description
Accelerometer
Gyroscope
Min
Typical
Max
Typical
Average
Two-line digital interface, standard I2C protocol
Value
±2g/ ±4g/ ±8g/ ±16g
±250/ ±500/ ±1000/ ±2000
2.375
3
3.46
500
2.7
Unit
g
°/sec
V
V
µA
µW
Table 3.4: Specifications of the Sensor MPU6050
d. Air quality sensor
The MQ135 is a widely used gas sensor designed for detecting a variety of gases in the air, including ammonia
(NH3), carbon dioxide (CO2), alcohol, benzene, and smoke. It is particularly popular in applications requiring air
quality monitoring, this sensor uses a metal oxide semiconductor (MOS) to detect the presence of gases and converts
this into an electrical signal.
Figure 3.7: The MQ135 sensor
• Sensor specifications:
– Gas Sensitivity:
∗ Ammonia (NH3 )
∗ Benzene (C6 H6 )
∗ Alcohol (ethanol, isopropanol)
∗ Carbon dioxide (CO2 )
∗ Sulfur dioxide (SO2 )
∗ Toluene (C7 H8 )
∗ Smoke
– Output: The MQ135 provides an analog output, which varies based on the concentration of gases.
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CHAPTER 3. PAYLOAD
– Power Supply: Operates at 5V, and it draws about 200mA during normal operation.
– Heating Element: The MQ135 has an internal heating element (at around 300°C) that helps to ionize
the air molecules, enhancing gas detection.
– Response Time: 10 to 30 seconds for detecting changes in gas concentration.
3.3.3
GPS module
The GPS NEO-8M is a popular GPS module developed by U-blox that provides accurate positioning data by receiving
signals from GPS satellites. It’s widely used in a variety of applications, from navigation systems to time synchronization, it supports both GPS (Global Positioning System) and GLONASS (Russian navigation system) satellite systems.
Figure 3.8: GPS module
• Module specifications:
– Positioning Accuracy: 2.5 meters horizontal and 5 meters vertical.
– Power Consumption: 50mA at 3.3V.
– Data Output: Outputs NMEA (National Marine Electronics Association) sentences over UART (Serial),
which include:
∗ GPGGA: Fix data (latitude, longitude, altitude, time).
∗ GPGLL: Geographic position.
∗ GPGSA: GNSS fix status.
∗ GPGSV: Satellites in view.
– Antenna: The module has a u.FL connector for an external antenna, improving signal reception and
accuracy.
3.3.4
SD Card Module
The microSD card module is a small, cost-effective board that interfaces with the microcontroller to provide access
to a microSD card. The module enables the reading and writing of data, typically through the SPI (Serial Peripheral
Interface) protocol.
Figure 3.9: SD Card Module
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CHAPTER 3. PAYLOAD
• Module specifications:
– Power Consumption: 20mA at 5V.
– Interface: Uses the SPI protocol to communicate with the microcontroller.
– Storage: The module supports microSD cards with a storage up to 32GB.
3.4
System Integration
For better results and to ensure reliable connections in the circuit with an optimal space distribution, we created a
customized printed circuit board to help achieve these goals, in this section, we are introducing the schematics of the
circuit followed by the PCB design.
3.4.1
Schematics
Using the KiCad software, we have developed the schematics that matches our desired circuit.
The idea here is to create a customized Arduino Mega shield that contains all the components of the circuit within
the interior of its surface to use minimum space, and is mounted right on top of the micro controller.
In the schematics you will see the different components and the wiring needed for every one of them with the pins
of the micro controller, a terminal block that is used to feed the circuit from the battery.
The figure below (Fig. 3.10) demonstrates the connections established in the circuit.
Figure 3.10: Payload schematics
This schematic of the payload circuit was first implemented in a prototype setup, as shown in Fig. 3.11 below. After
27
CHAPTER 3. PAYLOAD
verifying its functionality, the design was refined and prepared for PCB transformation, detailed in the next section:
PCB Design
Figure 3.11: Payload prototype circuit
3.4.2
PCB design
In order to create our PCB, we have placed all the components inside the shield in a suitable manner and started
establishing the connections and we can see the resulted circuit in the Fig. 3.12.
Figure 3.12: PCB
In the next figures (Fig. 3.12 and Fig 3.13) we see the 3D model that will help us visualize the real circuit.
28
CHAPTER 3. PAYLOAD
Figure 3.13: 3D Front
Figure 3.14: 3D Back
And here we present the final result of the printed circuit in Fig. 3.14 and Fig. 3.15.
Figure 3.15: Payload PCB first view
Figure 3.16: Payload PCB second view
Since we didn’t find a nearby PCB printer that can print both layers , we tried to create all the links in one layer and
we couldn’t avoid only two connections so we linked them manually using wires.
The final circuit with the sensors is presented in Fig. 3.17
29
CHAPTER 3. PAYLOAD
Figure 3.17: Payload circuit
3.5
Tests and validation
This section will treat the tests performed to ensure the correct performance of the payload.
3.5.1
Sensor calibration and tests
Before starting working on the circuit, we tested all the obtained sensors and listed below the detailed tests procedures.
30
CHAPTER 3. PAYLOAD
Sensor
BMP280
Tests
Table 3.5: Sensor Testing and Calibration
Calibration
• Pressure accuracy test: comparing against
a calibrated barometer under varying pressures.
• Temperature accuracy test: validating under
different controlled environments.
• Adjusting offsets for pressure and
temperature.
• Calibrating altitude formula using local sea-level pressure mentioned in
the data-sheet of the sensor.
• Altitude estimation: testing calculated altitude against known elevation levels.
AHT20
• Temperature dependence test: ensuring accurate humidity readings at various temperatures.
• Validating output against standard
humidity profiles.
• Drift test: measuring stability over extended
periods.
MQ135
• Response time test: measuring reaction
speed to gas concentration changes.
• Factor in environmental conditions
such as temperature and humidity.
• Baseline drift: monitor stability in clean air
over time.
MPU 6050
• Static accuracy test: verifying readings when
stationary.
• Aligning sensor orientation with
rocket axes.
• Dynamic accuracy test: validating against
controlled motion profiles.
• Drift test: monitoring long-term stability of
accelerometer and gyroscope.
N/A
General Testing
• Integration tests: ensuring that sensors work
properly with the Arduino Mega 2560 and
SD card.
• Environmental tests: validating performance
under vibration, pressure, and temperature
extremes.
3.5.2
PCB circuit continuity test
After done printing the circuit we perform a simple continuity test using a multimeter to ensure that there is no
discontinuities in the connections, and also the is no short circuits.
3.5.3
PCB stability test against vibrations and chock
During the flight the rocket is exposed to important vibrations that may affect the stability of our PCB.
CHAPTER 3. PAYLOAD
31
In order to test that , we will be using a vibration and chock generator, delivering amplitude bigger then the ones
encountered in the rocket.
The tests are scheduled for next week and will be shown in the presentation.
3.6
Conclusion
In this mission, our payload acts as a measurement unit, specially designed to gather environmental data during
the flight of the rocket to an altitude of 3 km. With this modest altitude, by no means did we stop ourselves from
engineering the payload as best as possible. By using high-precision sensors and housing the circuit in a customdesigned shield for the Arduino Mega 2560, we made the system robust enough to last under flight conditions.
This project showcases the team’s ability to work with constraints to deliver a practical, data-driven payload solution.
Although simple in scale, this mission lays the foundational stone to build up our experience in payload design and
deployment, opening ways for more ambitious missions.
Chapter 4
Recovery & Avionics
4.1
Introduction
In the early days of rocketry, launching a rocket was a one-way journey. Engineers watch their creation falls into
ruins after accomplishing their mission. Fortunately, a new mechanism has been developed to minimize the damage
taken by a rocket; it is the recovery system.
The primary aim of this mechanism is to slow down the descent of the aircraft, in order to be able to safely recover
the vehicle after flight. This step has been as crucial as the launch itself.
Moreover, this system integrates multiple components from mechanical to electrical ones, and specially designed
parachutes that must work in perfect synchronization to ensure successful recovery.
In this chapter, we will give an overview on the recovery system implemented in the rocket, followed by the integration of multiple electrical and mechanical hardware, processed using fine algorithms developed in specified software.
Lastly, we will present the tests assuring the reliability of the developed system.
4.2
System Overview
4.2.1
System’s description
A recovery system should be as simple and inexpensive as possible, highly reliable, lightweight and compact, and
should assure low drift of the rocket and final rate of descent not greater than 10 m/s.
In order to meet all these specifications, a dual-deployment recovery system is used. The first event occurs at apogee
(10 thousand feet) launching the main parachute, resulting in the separation of the rocket into two sections, the
nosecone, housing the payload, and the rocket tube that has the avionics bay (which also serves as a coupler). The
rocket then tumbles earthward, with the two sections connected by a tether. When 15 hundred feet altitude is reached,
the second event occurs. At this stage the remaining rocket body will be divided into two parts, the engine and the
avionics bay section. The two parts are combined to the drogue parachute with another line. These cables ensure
that the vehicle will remain intact during the whole flight, and will make finding all the parts after the landing easier.
The recovery stages are illustrated in the following figure (Fig 4.1).
The parachute deployment system utilizes a piston to isolate the parachute from the heat of the ejection charge and
to serve as a forward closure for the ejection charge compartment. The bulkhead of the avionics bay forms the
second closure of this last compartment. The force acting on the piston, when the charge fires, blows the upped part
(nosecone housing the payload) out the rocket body. The piston is pulled out at the same time, being attached to the
tether connecting the avionics bay to the upper body.
This dual-deployment system presents several advantages over the single event one. As it incenses the reliability of a
safe landing, the first event separates the rocket into two pieces, generating an aerodynamically unstable vehicle and
therefore relatively slow descent.
This descent is still relatively quick, as it is fast enough to reduce the overall descent time. As such, wind has less time
32
CHAPTER 4. RECOVERY & AVIONICS
33
Figure 4.1: Dual deployment stages
to carry the rocket downrange.
The system is monitored and controlled by a flight computer, which functions as both a data acquisition system and
a launch control unit. This micro-controller makes the decision based on the altitude data gathered from a pressure
sensor, and is transferred to a ground station in real time, to deploy the parachute manually in case of automatic
deployment failure, serving as a backup system and increasing redundancy.
4.2.2
Avionics block diagram
This process is controlled by a compact circuit comprising a micro-controller and various other electrical components, which are interconnected and interact as follows:
Figure 4.2: Avionics block diagram
In order to detect the apogee with accuracy and launch the parachute effectively, a barometric pressure sensor measures the atmospheric pressure that changes in different altitudes. This signal is then transferred to the micro-
CHAPTER 4. RECOVERY & AVIONICS
34
controller, where it processes the data and deduce the altitude. According to these measurement and calculations, the
ejection charge can be activated at the apogee or the second event. A command signal is then sent to the fire charge,
which starts a small explosion leading to the deployment of the parachute.
The data is sent in real time to a ground station, where it will be supervised and visualized, in case of any failure of
the deployment, the fire charge can be activated manually, using a transmission module.
4.3
Hardware components
4.3.1
Flight computer
To guarantee a safe flight and successful landing, a reliable flight computer is essential for controlling parachute
deployment based on altitude readings. For this purpose, we selected the Heltec WiFi LoRa 32 (V2) board.
The Heltec LoRa32 V2 is a compact and versatile microcontroller board designed for IoT applications, featuring LoRa
(Long Range) communication capabilities. It combines a powerful ESP32 dual-core processor with integrated Wi-Fi
and Bluetooth, making it suitable for projects requiring connectivity and long-range wireless communication.
Figure 4.3: Heltec LoRa32 V2 micro-controller
• Board technical specification
CHAPTER 4. RECOVERY & AVIONICS
Parameters
Master Chip
LoRa Node Chip
USB to Serial Chip
Frequency
Wi-Fi
Bluetooth
Hardware Resource
Memory
Interface
Battery
Operating Temperature
Dimensions
35
Description
ESP32 (240MHz Tensilica LX6 dual-core+1 ULP, 600 DMIPS)
SX1276/SX1278
CP2102
470 510MHz
802.11 b/g/n, up to 150Mbps
Bluetooth V4.2 BR/EDR and Bluetooth LE specification
22*GPIO;
3*UART;
2*I2C;
2*SPI;
520KB internal SRAM;
8MB Spi Flash
Micro USB x 1;
LoRa ANT(IPEX1.0);
2*18*2.54 Header Pin
3.7V lithium (SH1.25 x 2 socket)
-20 70 ℃
51 * 25.5 * 10.6 mm
Table 4.1: Specifications of the Device
4.3.2
Ground station
At the ground station, a similar LoRa32 V2 board is employed to establish communication with the onboard system.
This board is responsible for receiving real-time flight data transmitted via LoRa and providing control over the
backup deployment system, ensuring a secondary layer of safety and reliability during the mission.
4.3.3
Sensors
Since the parachute deployment is directly linked to the rocket’s altitude, a BMP280 barometric pressure sensor is
employed to provide continuous updates on altitude. This ensures precise monitoring and accurate triggering of
the deployment mechanism at the appropriate height. (the technical specifications for the sensor are mentioned in
chapter 3).
4.3.4
Deployment trigger circuit
a. MOSFET
The IRFZ44N N-channel MOSFET (Metal-Oxide-Semiconductor Field-Effect Transistor) is utilized to drive the ejection charge. With a continuous drain current capacity of up to 49A and a gate voltage requirement of 10V, it offers a
safe and reliable solution for this critical function.
Figure 4.4: The IRFZ44N MOSFET
b. Optocoupler
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CHAPTER 4. RECOVERY & AVIONICS
To isolate the control circuit from the power circuit, we implemented a TPL250 optocoupler. This component allows
the microcontroller to safely and efficiently drive the MOSFET gate by providing electrical isolation, ensuring reliable
operation and protecting the control electronics from potential high-power surges.
Figure 4.5: Optocoupler
c. Ejection charge circuits (E-matches)
E-matches (electronic matches) are small, electrically ignitable devices used in pyrotechnics and explosive applications. They consist of a heated bridge wire surrounded by a pyrotechnic composition. When an electrical current is
passed through the bridge wire, it heats up and ignites the pyrotechnic material, triggering the desired effect.
Figure 4.6: E-matches
Due to the unavailability of commercially-sourced e-matches, we took the initiative to fabricate our own by utilizing
electrical wires and standard matches. This involved adapting readily available components to meet the necessary
specifications for reliable ignition in our system.
4.4
System Integration
Having outlined the hardware components, the focus moves to the system’s wiring and interconnections, detailing
how these components are integrated to function cohesively. Tacking in consideration all the noise factor that may
interrupt the circuit.
4.4.1
Schematics
To ensure the reliable deployment of the recovery system in rocketry, the wiring diagram in Fig. outlines the electrical
connections and integration of key components, including the flight computer, power supply, deployment charges,
and sensors.
Before all, a power supply of 12 volts is lowered down to 5 volts using a buck converter, to supply the EPS 32 V2
microcontroller.
From the microcontroller part, it powers the BMP280 sensor with 3.3 volts, then gathers the data through the I2C
communication pins (SDA & SCL).
After processing and making a decision based on the information it has, the ESP 32 will deploy the parachute at the
right time by sending a command signal to the MOSFET so it can fully open its gate and let a huge current flow from
CHAPTER 4. RECOVERY & AVIONICS
37
Figure 4.7: Deployment system Schematics
the drain to the source. And therefore, firing the electrical matches.
However, since the ESP 32 can only provide 3.3 volts at the output of each pin, an optocoupler is needed to amplify
that signal before feeding it to the MOSFET.
In addition, the ESP32 V2, equipped with a LoRa module, enables real-time altitude monitoring.
This schematic (Fig. 4.7) represents the blueprint of the flight board circuit, which was initially implemented as a
prototype in real life. Together with the receiver at the ground station (Fig. 4.8), this setup forms a redundancy
system, ensuring the security and reliability of the deployment circuit. Extensive testing and validation of this circuit
confirmed its functionality, paving the way for the next stage: transforming it into a printed circuit board (PCB).
Figure 4.8: Deployment system circuit: transmitter and receiver
CHAPTER 4. RECOVERY & AVIONICS
4.4.2
38
PCB design
To ensure that all the electrical components stay intact in their places, a PCB is the best option. The following figure
present the PCB designed for the previous circuit:
Figure 4.9: PCB Design for the deployment system circuit
To help vision the real circuit, we present the 3D model in Fig. 4.10.
Figure 4.10: 3D representation of the PCB for the deployment system circuit
And finally, the real printed circuit in Fig 4.11.
4.4.3
EMI considerations
Electromagnetic Interference EMI can disrupt the desired operation of electronic circuits by introducing unwanted
noise and signals, which is particularly critical in sensitive systems like this one.
To mitigate these issues, several measures have been implemented in the design. Power and signal wires are routed
CHAPTER 4. RECOVERY & AVIONICS
39
Figure 4.11: PCB of the deployment system circuit
separately to minimize the risk of coupling interference, ensuring that high-power circuits do not affect the lowpower signal paths. Additionally, a single-point grounding system is used to prevent ground loops, which could
introduce noise and lead to erratic behavior in the system.
4.4.4
Isolation
Another factor in the stability of performances of the circuit rely on isolating the power and command circuits. By
using an optocoupler, we prevent direct electrical contact between these circuits, which reduces the risk of voltage
spikes or noise from the power circuit affecting the control system.
4.5
Software
With the hardware components and wiring established, we now shift our focus to the software framework that
governs the system’s functionality and ensures seamless deployment operation.
4.5.1
Flight states/phases
The vehicle flight goes through 3 stages as mentioned in the system’s overview.
In this initial phase, the rocket is propelled upward by the engines, with the flight computer continuously monitoring
key parameters such as altitude, and pressure. During ascent, the primary objective is to reach the desired altitude of
10,000 feet while ensuring the structural integrity of the rocket.
Upon reaching the apogee, the rocket will begin to coast. At this point the engine goes off, and the rocket begins free
fall. The rate of descent as well as the altitude of the aircraft is continuously monitored by the onboard computer.
The flight computer receives new sensor information during the entire flight. This stage is important to determine
if the conditions for the deployment of the recovery system have been satisfied.
After reaching the apogee, the descent phase begins. The flight computer monitors the altitude to determine the
appropriate moment for deploying the main parachute. The recovery phase ends when the rocket safely returns to
the ground, where it can be retrieved for post-flight analysis.
CHAPTER 4. RECOVERY & AVIONICS
4.5.2
40
Altitude calculation method
By utilizing the BMP280 sensor, pressure data is fed to the ESP32, enabling the calculation of altitude. This is achieved
through the following Hypsometric equation:
!
1
5.255
T + 273.15
P0
(4.1)
Altitude =
× 1−
0.0065
P
Where:
• T : Temperature at the current location measured by BMP.
• P : Pressure at the current location measured by BMP.
• P0 : Baseline pressure.
Consequently, the relationship between altitude and pressure variation is illustrated in the following plot in Fig 4.12.
Figure 4.12: The relationship between altitude and pressure
This method, however, has certain limitations. The primary issue lies in pressure variations; at a given altitude, the
pressure can fluctuate, leading to inaccuracies in the calculated altitude.
To address this issue, we employ a Kalman filter, which helps to reduce the impact of pressure fluctuations and
improve the accuracy of altitude estimation.
4.5.3
Deployment algorithm
In this section, we present the logic and steps implemented in deployment algorithm, which processes real-time
sensor data to determine the optimal conditions for triggering the parachute deployment.
The deployment algorithm is presented in the following diagram in Fig 4.13.
For each event, we outline the conditions that must be met before initiating the deployment algorithm. In the first
event, the altitude must exceed 2700 m to begin monitoring and awaiting the apogee. For the second event, the first
event must have been completed (indicated by the Boolean variable main_parachute being true, which confirms the
deployment of the main parachute), and the altitude should be below 650 m.
4.6
Testing and Validation
4.6.1
Sensor calibration and tests
To evaluate the BMP280 sensor, we designed a series of tests divided into four parts. First, we assessed its precision
by measuring small variations in altitude of 10 cm (test 1) and 20 cm (test 4). Next, we tested its responsiveness to
sudden altitude changes of 1 meter (test 2). Finally, we examined its ability to maintain accuracy over an extended
period while remaining stationary at the same position (test 3).
The tests take time from:
CHAPTER 4. RECOVERY & AVIONICS
41
Figure 4.13: Deployment algorithm
• Test 1: t1 = [0, 48Ts ]
• Test 2: t2 = [48Ts , 56Ts ]
• Test 3: t3 = [56Ts , 110Ts ]
• Test 4: t4 = [110Ts , 142Ts ]
Where: Ts is the sampling time of the micro-controller, Ts = 1.1 s.
We visualize the performances of the sensor in the graph in Fig 4.14.
The altitude measurement showed a reasonable tracking of the reference, but exhibited some fluctuations, indicating
CHAPTER 4. RECOVERY & AVIONICS
42
Figure 4.14: Measured altitude from sea level
the need for a filter to improve stability and accuracy.
So, for the next tests we introduce a Kalman filter, the results are shown in the following graph 4.15.
Figure 4.15: Filtered measured altitude
After implementing a Kalman filter, the fluctuations in the altitude measurements decreased, but they are still present
to some extent.
Overall, the BMP280 sensor performed within acceptable accuracy limits, making it suitable for integration in the
deployment circuit.
4.6.2
Deployment circuit test
In this phase of the tests, we simulated the deployment process on a small scale. To do this, we sent the flight computer
up to an altitude of 2.5 meters and then brought it back down. During the test, we monitored the altitude data and
counter from the computer’s serial monitor, and recorded the point at which the parachute was virtually deployed,
indicated by the flight computer sending a message confirming the deployment.
In a real rocket scenario, the sampling time would be set to 0.5 seconds, as the micro-controller needs to detect the
descent at least three times before initiating parachute deployment. To ensure accurate detection within an altitude
range of 3,000 m ± 150 m (the specified error margin), we calculated the time required for the descent to pass 150
m. We found that it would take approximately 9 seconds to descend this distance, meaning the apogee needs to
be detected within this timeframe. With this in mind, we concluded that a sampling time of 0.5 seconds would be
CHAPTER 4. RECOVERY & AVIONICS
43
optimal for the real rocket, allowing for around 1.5 seconds to detect the apogee.
Using the same logic, we attempted to detect the apogee in the small-scale test with a 0.5-second sampling time to
ensure the system would function properly on a larger scale. So, for Ts = 0.5 s, we get the altitude diagram in Fig
4.16.
Figure 4.16: Apogee detection test results
The altitude diagram obtained with a sampling time of 0.5 seconds showed that the apogee was detected at t = 25·Ts
which is an acceptable result.
4.7
Conslusion
This chapter presented an overview of the rocket’s recovery system, detailing the integration of various hardware
components, including mechanical and electrical elements. It also discussed the algorithms designed to control these
components for effective recovery. The chapter concluded by emphasizing the thorough testing conducted to validate
the system’s reliability, ensuring that the recovery mechanism functions as intended and is fully prepared for future
missions.
Chapter 5
Structural Strength Analysis
Having detailed the dimensioning of each component, we now turn our attention to ensuring these dimensions meet
the necessary strength and performance criteria.
5.1
Main Rocket
5.1.1
Fin Flutter
Fin flutter occurs when a model rocket’s fins experience aerodynamic forces that cause them to vibrate at high frequencies. This can be triggered by a combination of factors, including speed, fin size and shape, and material stiffness.
The problematic is when the rocket speed exceeds the maximum fin flutter speed at which point the air will amplify
oscillations to the point of destroying the fin. The maximum fin flutter can be calculated from the following formula:
v
u
G
Vf = au
(5.1)
t 1.337×(AR)3 ×P ×(λ+1) (Howard, 2011, 3)
3
2×(AR+2)×( Ctr )
AR =
b2
(Fin Semi-span length or height)2
=
Fin Area
S
S=
(dimensionless)
1
(Cr + Ct ) × b
2
(5.2)
(5.3)
• Ctr : The fin thickness ratio
t
Fin Thickness
=
Cr
Root Chord Length
(dimensionless)
• λ: The fin taper ratio
λ=
Ct
Tip Chord Length
=
Cr
Root Chord Length
(dimensionless)
• P : Air pressure at the altitude where the speed of sound was determined (Pa) (Atmospheric Pressure Vs. Elevation Above Sea Level)
5.26
P = P0 × 1 − 2.5577 × 10−5 × h
• P0 : Air pressure at sea level,
P0 = 101325 Pa
• h: Altitude where maximum velocity was reached (m)
Numerical Application for the main rocket:
• Mechanical properties our composite material:
44
45
CHAPTER 5. STRUCTURAL STRENGTH ANALYSIS
Density ρ (g/cm3 )
1.5
Young Modulus E (GPa)
59
Shear Modulus G (GPa)
33
Table 5.1: Mechanical properties of the composite material
• At h = 847 m: V = Vmax = 278.3 m/s, hence a = 337.05 m/s
2
2
22
• S = 12 × (29 + 9) × 22 = 418 cm2 , hence AR = bS = 296
= 1.158
9
= 0.31
• λ = 29
• Ctr = 0.3
29 = 0.01
• P = 101325 × 1 − 2.5577 × 10−5 × 847
5.26
= 90298.42 Pa
v
u
u
33 × 109
Vf = 337.05 × t 1.337×(0.864)3 ×90298.42×(0.31+1)
2×(1.158+2)×(0.01)3
Vf = 297.6 m/s
This velocity we calculated is the critical speed at which rocket’s fins are likely to experience "flutter", since our rocket’s
maximum velocity is 279 m/s, our rocket is candidate to resist fin flutter effectively.
Final thoughts
• The selection of “G” itself has a fair bit of windage. Virtually none of the materials in common use for amateur
rocket fins come with guaranteed manufacturing specifications. This is especially true for composite materials.
• The equations as presented are only valid in the Troposphere (under 36,000 ft./10,973 m or so).
• Vf increases as we gain altitude because air temperature and air pressure decrease. Unless we are well into the
supersonic range, fin flutter is not likely to be a problem at high altitude.
• This method for fin analysis does not address the need for strong fin attachment, but any attachment method
will benefit from not being stressed by fin flutter. (Bennett, 2023)
5.1.2
Aerodynamic Loading
The equation for rocket aerodynamic drag was given earlier by Equation 1:
Fd =
1 2
ρv Cd A
2
Numerical Application for the main rocket:
• Maximum velocity: Vmax = 279 m/s
• Drag Coefficient: Cd = 0.475 (retrieved from OpenRocket)
• The outer diameter: D = 0.194 m ⇒ A = 0.0295 m2
• From Refprop, the air density: ρ = 1.1348 kg/m3
(5.4)
CHAPTER 5. STRUCTURAL STRENGTH ANALYSIS
46
Figure 5.1: REFPROP properties at maximum velocity
1
× 1.1348 × (279)2 × 0.475 × 0.0295 = 618.9 N
2
As seen, drag force Fd acting against a rocket’s upward motion is proportional to the body diameter squared, meaning
doubling the diameter (D) of the rocket body increases drag force by a factor of four. Or put another way, reducing
the body diameter by ½ reduces drag losses to ¼. Also, drag force increases dramatically as velocity increases, so due
diligence is wise.
Fd =
5.1.3
Flight loads
Another factor to consider with regard to structural analysis is strength. In flight, a rocket body is subjected to
compression and bending loads. Bending load is a consequence of non-zero angle of attack of the flight path.
• Angle of attack
Gusts, acting horizontally, generate normal forces on the rocket body. The term normal refers to loads that are
applied perpendicular to the rocket’s long axis, or side loads. A sudden gust velocity vector (U), when added to the
vehicle’s velocity vector (V) at end of the launcher, causes a small angle of attack alpha to the rocket’s flight path
(Fig.33).
Figure 5.2: Angle of Attack
Hence: α = tan−1
U
V
According to the National Association of Rocketry, launching is not advisable when wind speeds exceed 8 m/s (20
mph) (Maximum Wind Speeds for Launching a Model Rocket, 2023).
And a rule of thumb states that an angle of attack should not exceed 10°, based on that we found that:
In moderate conditions (Leaves & small twigs move, light flags extend) U=3.5-5 m/s (The University of Maine.), we
take:
U = 4 m/s
47
CHAPTER 5. STRUCTURAL STRENGTH ANALYSIS
This angle of attack causes lift forces on the nose and fins (transitions and boat-tail if present).
As mentioned before, flying at some non-zero angle of attack, whether caused by gust or dynamic imbalance generates
normal forces of the rocket acting on the nosecone, fins, transitions and boattail (if present). In order for these forces
to develop, something has to react these forces. This something is the inertia of the rocket, both transverse inertia
and rotational inertia. The resulting forces generate a bending moment being applied to the rocket body. (Richard
Nakka’s Experimental Rocketry Site, 2022)
a. Lateral Shear and Bending Moments
The normal force acting upon a particular rocket component such as rocket body, nosecone, conical transition (such
as shoulder or boattail) and fins due to non-zero angle of attack is calculated as such:
Nx = q × A × α × ((CN )α )x
(5.5)
Where:
• q = dynamic pressure = 21 ρV 2 , N/m2
• A = reference area = 14 πD2 , where D = nosecone base diameter, m.
• V = vehicle velocity, m/s
• ((CN )α )x = slope of the normal force coefficient at α = 0 ( ∂CN
∂α α=0 ), per radian.
It can either be retrieved from OpenRocket or calculated using the Barrowman Method. (See Appendix A)
• α = effective angle of attack, radians
• XCG , XCP = center of gravity and center of pressure, retrievable from OpenRocket.
• XN , XF = locations of the centers of pressure for the nose cone and the fins, respectively.
X1 and X2 may be calculated as:
X1 = XCG − XN
X2 = XF − XCG
The values of distributed inertia loads w1 and w2 are next calculated:
W1 =
W2 =
NN + NF − W2 X2
X1
NF (2X2 + X1 ) − NN X1
X22 + X1 X2
The value of w2 is calculated first, then w1 .
Unit of w1 and w2 is N/mm2 .
The lateral shear V as a function of x, where x is the distance along the rocket body with x = 0 at nose CP, is given
by:
V (x) = NN − W1 x for 0 < x < X1
V (x) = V1 − W2 (x − X1 ) for X1 < x < L
L represents the body length from nose CP to the fin CP: L = X1 + X2 .
The bending moment M as a function of x is given by:
W1 x2
for 0 < x < X1
2
1 2 1 2
M2 (x) = V1 x − W2 X1 x + L − x − L(V1 + W2 X1 ) for X1 < x < L
2
2
M1 (x) = NN x −
Units: Newtons and millimeters. First, the angle of attack:
Voff-rod = 25.2 m/s
48
CHAPTER 5. STRUCTURAL STRENGTH ANALYSIS
α = tan−1
U = 4 m/s
4
= 9.02◦ = 0.1570 rad
25.2
α = 0.1570 rad.
Reference diameter: D = 194 mm
Reference area: A = 295.6 × 10−4 m2
Velocity:
V = Vmax = 279 m/s
Air density where V = Vmax :
ρ = 1.1348 kg/m3
The dynamic pressure is therefore:
1
× 1.1348 × 2792 = 46411.05 N/m2 = 0.04641 N/mm2
2
The normal force acting on each component is now calculated:
q=
• Nose cone
Nosecone profile is tangent ogive.
(CNα )N = 2
NN = 0.05173 × 29560 × 0.1396 × 2 = 272.6 N
Applied at center of pressure: XN = 0.466L
XN = 0.46679 = 36.814cm
• Body tube
For any rocket body, (CN )α = 0, hence NB = 0 N
• Fins
(CNα )F = 2
NF = 0.044166 × 29560 × 0.1570 × 7.776 = 1593.85 N
Applied at center of pressure:
XF = XB +
XR
3
Cr + 2Ct
Cr + Ct
+
1
6
2Cr Ct
Cr + Ct −
Cr + Ct
• XB = L − Cr = 217 cm
• Root cord, Cr = 28 cm
• Tip cord, Ct = 8.12 cm
• XR = Cr − Ct = 19.88 cm
• Hence,
• XF = 230.087 cm
• X1 = XCG − XN = 148 − 36.81 = 111.19 cm
• X2 = XF − XCG = 230.09 − 148 = 82.1 cm
• L = 111.19 + 82.1 = 193.29 cm
The lateral acting inertia loads, w1 and w2 , as shown in the figure below, are now calculated:
W2 =
NF (2X2 + X1 ) − NN X1
1593.85 (2 × 82.1 + 111.19) − 272.6 × 111.19
=
X22 + X1 X2
(82.1)2 + 82.1 × 111.19
W2 = 2.58 N/mm
W1 =
NN + NF − W2 X2
272.6 + 1593.85 − 2.58 × 82.1
=
= −0.226 N/mm
X1
111.19
49
CHAPTER 5. STRUCTURAL STRENGTH ANALYSIS
Figure 5.3: Aero and inertia loads acting on rocket body in flight
Inserting the values in the lateral shear expressions:
V (x) = 272.6 + 2.21x for
V1 = V (X1 ) = 272.6 − 2.58X1
⇒
0 < x < X1
V1 = 272.6 + 0.226 × 1111.9 = 523.9 N
V (x) = 523.9 − 2.58(x − 1111.9) for
X1 < x < L
VL = V (L) = 523.9 − 2.58(L − 1111.9) = 523.9 − 2.58 × (1932.9 − 1111.9) = −1594 N
The maximum bending moment occurs where V = 0. Since this location may be anywhere along x, both equations
for V (x) are solved in terms of x.
272.6
NN
=
= −123.35 mm
W1
−2.21
As this value of x is outside the range of x = 0 to X1 = 1111.9 mm, the true location of V = 0 must lie in the range
x = 1111.9 mm to x = L = 1932.9 mm, and therefore,
x(V =0) =
V1
523.9
+ X1 =
+ 1111.9 = 1315.96 mm
W2
2.58
The maximum bending moment is therefore given by:
1 2 1 2
V12
− V1 (L − X1 ) + W2
L + X1 − LX1
Mmax =
2W2
2
2
x(V =0) =
50
CHAPTER 5. STRUCTURAL STRENGTH ANALYSIS
Plugging in the values:
Mmax =
523.92
− 523.9 × (1932.9 − 1111.9) − 0.226 ×
2 × 2.58
1
1
(1932.9)2 + (1111.9)2 − 1932.9 × 1111.9
2
2
Mmax = −453.1 N.m
The bending moment and shear stress distributions indicate the points along the rocket body where the maximum
stresses occur. These points are critical for ensuring the structural integrity of the rocket.
Figure 5.4: Effects of Shear and Bending Moment on rocket
Now that we have a suitable means of representing shear and bending moment as a function of x along the rocket
body, we can generate Shear and Bending Moment diagrams
CHAPTER 5. STRUCTURAL STRENGTH ANALYSIS
51
Figure 5.5: Shear force Vs Distance
Figure 5.6: Bending moment Vs Distance
b. Interpretation of the Results Obtained
1. Shear Force Distribution
• It can be seen that the shear force is increasing within the first segment (From nose tip to X1 ), which indicates
that the upper body section experiences an increasing load as we move towards X1 .
• After X1 , the shear force decreases linearly, reflecting the change in loading conditions as you move towards
the aft end of the rocket. This behavior is expected, as the lateral inertia loads act in opposing directions.
(Figure 17)
2. Bending Moment Distribution
52
CHAPTER 5. STRUCTURAL STRENGTH ANALYSIS
• Similarly to the lateral shear, in the first segment, the bending moment is increasing, but not linearly, reaching
a peak value before transitioning at X1 , where a sudden drop can be seen, implying that the middle section of
the rocket is subject to the highest bending stresses.
• After X1 , the bending moment decreases steadily. This suggests that the loads acting on the rocket cause a
bending moment that diminishes towards the end of the rocket.
3. Critical Points
• The transition at X1 (111.19 cm) is a critical point where both the shear force and bending moment exhibit
changes in behavior. This indicates a potential weak point in the rocket’s structure. Therefore, this area should
be carefully inspected and reinforced to prevent structural failure during launch and flight.
5.1.4
Calculating parachute size and resistance
For a hexagonal shaped parachute such as the one we used, the diameter is:
√
D = 0.866 × S
Numerical application:
• The mass of the rocket with empty engine, 22,387 kg.
• The gravity, g = 9.81 m/s2 .
• Drag coefficient, Cd = 0.8.
• Air density where the h = 457 m is ρ = 1.17 kg/m3 .
• Air density where the h = 3071 m is ρ = 0.91 kg/m3 .
• Descent velocity for the drogue parachute, V = 16.4 m/s.
• Descent velocity for the main parachute, V = 7.2 m/s.
Smain = 9.05 m2
Sdrogue = 2.24 m2
p
0.866 × Smain = 279 cm
p
Ddrogue = 0.866 × Sdrogue = 140 cm
Dmain =
A parachute must be dimensioned to resist opening. The shock on opening can be assessed as follows
F =
ρSCd V02
2
Since the drogue parachute is deployed at the apogee at nearly zero speed, its cords are not subjected to significant
stresses. Therefore, the calculation is only relevant for the main parachute, which is deployed at a much higher speed,
resulting in considerably greater forces on its suspension lines.
• Drag coefficient, Cd = 0.8
• Air density at h = 457 m is ρ = 1.17 kg/m3
• Opening velocity V0 = 16.4 m/s
• Area of the main parachute, Smain = 9.05 m2
• Number of suspension lines for the main parachute, Nmain = 12 lines
53
CHAPTER 5. STRUCTURAL STRENGTH ANALYSIS
For the main parachute: Fmain =
ρSCd V02
= 1139.15 N
2
The force calculated in this way applies to the parachute’s attachment point on the spindle, to the webbing, to the
swivel, to all the lines and to the parachute itself. For the lines, don’t divide the force obtained by the number of lines,
as their different lengths render some of them useless at the moment of opening. Consider, that only 75% of the lines
contribute to opening resistance.
0.75 × Nmain = 9 lines
Fsm =
1139.15
Fmain
=
= 126.57 N
0.75 × Nmain
9
The length of the suspension lines is calculated using the formula:
L = 1.5 × D
For the main parachute:
L = 1.5 × 300 = 450 cm
For the drogue parachute:
5.2
L = 1.5 × 150 = 225 cm
Structural Strength Calculations Low Altitude Rocket
a. Drag force
Fd =
1 2
ρV Cd A
2
(N.A) :
• Maximum velocity: Vmax = 295 m/s
• Density: ρ = 1.1768 kg/m3
• Drag Coefficient: Cd = 0.598, retrieved from OpenRocket.
• The outer diameter: D = 1.1 m ⇒ A = 0.0095 m2
Fd =
1
× 1.1768 × (295)2 × 0.598 × 0.0095 = 290.9 N
2
b. Angle of attack
−1
α = tan
U
V
(N.A) :
• U = 4 m/s
• Voff-rod = 23.5 m/s
−1
α = tan
4
23.5
= 9.78◦ = 0.157 rad
c. Normal force
Nx = q × A × α × (CNα )x
(N.A):
• α = 0.157 rad.
• Reference diameter, D = 110 mm
54
CHAPTER 5. STRUCTURAL STRENGTH ANALYSIS
• Reference area, A = 9500 mm2
• V = Vmax = 295 m/s
• Air density where the V = Vmax is ρ = 1.1768 kg/m3
The dynamic pressure is therefore:
q=
1
×ρ×V2
2
1
× 1.1768 × 2952 = 51205.51 N/m2 = 0.051205 N/mm2
2
(CNα )x is retrieved from Openrocket.
q=
The normal force acting at each component is now calculated:
• Nose cone
Nosecone profile is tangent ogive.
(CNα )N = 2
NN = 0.051205 × 9500 × 0.157 × 2 = 152.74 N
Applied at center of pressure: XN = 0.466L
XN = 0.466 × 33 = 15.37cm
• Body tube
For any rocket body, (CNα )B = 0, hence NB = 0 N
• Fins
(CNα )F = 9.673
NF = 0.051205 × 9500 × 0.157 × 9.673 = 738.74 N
Applied at center of pressure:
XR
XF = XB +
3
Cr + 2Ct
Cr + Ct
1
+
6
2Cr Ct
Cr + Ct −
Cr + Ct
• XB = L − Cr = 109 cm
• Root chord, Cr = 14 cm
• Tip chord, Ct = 12 cm
• XR = Cr − Ct = 2 cm
• XF = 113.23 cm
• Parachute calculations
The area of the parachute from the following equation:
2mg
S=
ρCd V 2
√
D = 0.866 × S
(N.A) :
• The mass of the rocket with empty engine, m = 2.372 kg
• The gravity, g = 9.81 m/s2
• Drag coefficient, Cd = 0.8
• Air density where the altitude h = 1539 m is ρ = 1.05 kg/m3
• Descent velocity, V = 7.71 m/s
S = 0.9177 m2
√
D = 0.866 × S = 103 cm
Chapter 6
Softwares: Analysis and Simulations
We now have a general idea of what our rocket will look like, as outlined in Chapter 2. It is time to simulate our
design choices to verify their validity. We used various software tools, each tailored to specific areas of analysis and
simulation.
6.1
Openrocket
Openrocket was used extensively as the primary modeling software of the rocket and its flight simulations.
All performance metrics as stated in this paper are gathered from Openrocket unless otherwise stated.
We designed our rocket to be aerodynamically efficient, lightweight, and stable while achieving its target altitude,
all while meeting the competition requirements. After experimenting with various designs and analyzing multiple
calculations, we have concluded that the following design is our optimal choice. (Figure)
Figure 6.1: Skybridge design in Openrocket
Every small change was followed by a simulation in OpenRocket to ensure everything was in place.
55
CHAPTER 6. SOFTWARES: ANALYSIS AND SIMULATIONS
56
Figure 6.2: Skybridge design in Openrocket
6.1.1
Thrust curve
The graph shows the thrust curve for the Cesaroni 10367N1800-P (Figure 4). The y-axis represents the thrust in
Newtons, going from 0 to 2250 N, and the x-axis shows the time in seconds, from 0 to 6 seconds. When the motor
starts, it produces around 2000 N of thrust, quickly rising to 2206 N, which is the maximum thrust it reaches. After
that, the thrust gradually drops off as the motor burns fuel, eventually reaching 0 N at the 6-second (Burn out).
Figure 6.3: Thrust curve
6.2
Rocksim
Rocksim, a software similar to OpenRocket was used during the process of building our rocket to ensure that we
were on the right track. As shown in the figure below, the results obtained for weight, center of gravity, center of
pressure, and stability are almost identical, further validating our design.
Figure 6.4: Final design Rocksim
CHAPTER 6. SOFTWARES: ANALYSIS AND SIMULATIONS
57
Similar to the procedure used in OpenRocket, various simulations were conducted and repeated whenever a small
change occurred. Fortunately, all of them validated our design choice.
Figure 6.5: Simulations on Rocksim
6.2.1
Thrust curve
Figure 6.6: Thrust curve
6.3
Openrocket Vs Rocksim
The values we obtained from both softwares are very close (see similarities).
However, there are some minor differences, the main reason for this is that OpenRocket’s motor differs from RockSim’s which can influence the rocket’s trajectory, apogee, thrust curves, and overall performance. (See differences)
6.3.1
Similarities
The stability margins in the figure and figure are nearly identical, with a stability of 2.2 Cal in OpenRocket and 2.3 in
RockSim.
The same figures indicate that the centers of gravity and pressure are also similar. OpenRocket records the center
of gravity at 148 cm and the center of pressure at 190 cm, while RockSim reports them as 145 cm and 190 cm,
respectively.
Additionally, the acceleration values and velocity off rod are the same in both software programs.
CHAPTER 6. SOFTWARES: ANALYSIS AND SIMULATIONS
58
Finally, the figures below demonstrate that the trajectories of both simulations are identical, showing that the times
for burnout and apogee match perfectly.
Figure 6.7: Altitude Vs Time or (Trajectory of flight)
6.3.2
Differencies
The analysis indicates that, for the same motor (Cesaroni 10367N1800-P) used in both OpenRocket and RockSim,
there were no significant differences in core specifications such as burnout time, deployment velocity, and maximum
acceleration. However, a notable discrepancy was observed in the apogee, which was higher in RockSim (3507 meters) compared to OpenRocket (3010 meters).
Although both tools rely on similar simulation algorithms, variations in aerodynamic or environmental assumptions—such as drag coefficients, wind profiles, or the resolution of trajectory modeling—may have influenced this
outcome.
Furthermore, as the motor used was not available in RockSim’s original library, it was manually added to the software’s database. While every effort was made to ensure the accuracy of the input data, this process might have
contributed to the observed variation in apogee.
CHAPTER 6. SOFTWARES: ANALYSIS AND SIMULATIONS
6.4
59
SOLIDWORKS
SOLIDWORKS was used to evaluate the structural integrity of some our rocket components under various load
conditions. This analysis will help us understand how the parts will respond to static forces and identify any potential
weaknesses.
Figure 6.8: Units used in static simulation
6.4.1
Simulation of the nosecone
In the static study simulation, the nosecone was constrained at its rear (indicated in green) to simulate the fixed
attachment point. A parallel force, representing the air drag encountered during flight, was applied uniformly over
the rocket’s surface to simulate aerodynamic forces acting on the body (indicated in purple).
Figure 6.9: model nosecone
The table below, shows the different properties of the nosecone :
CHAPTER 6. SOFTWARES: ANALYSIS AND SIMULATIONS
60
Figure 6.10: Different properties of the nosecone
6.4.2
Loads and Fixtures
When the load named “Force-1” representing the aerodynamic forces acting on the rocket during flight is applied, a
reaction force is expected in direction X, Y, and Z. The resultant force is the net effect of all the forces acting on the
rocket. In our case, a force of 700 N in the direction of flight (presumably along the nose cone of the rocket).
Figure 6.11: Loads and Fixtures
6.4.3
Study Result
a. Stress Distribution
Von Mises stress is a yield criterion used to predict when materials will yield under complex loading. It combines
the effects of normal and shear stresses into a single value.
Low Von Mises stress areas indicate regions where the material is under less load. In our simulation, the minimum
Von Mises stress is approximately 2.4 × 10−7 N/m2 .
High Von Mises stress areas suggest potential failure points, where the maximum Von Mises stress is 1.8×105 N/m2 .
If the Von Mises stress exceeds the material’s yield strength, it may deform or fail. In our case, the maximum stress
CHAPTER 6. SOFTWARES: ANALYSIS AND SIMULATIONS
61
is well below the material’s yield strength of 7 × 107 N/m2 , indicating that our nose will not yield or fail under the
applied loading conditions.
Stress :
Figure 6.12: nosecone stress
b. Displacement
Resultant Displacement
It indicates how much the nosecone moves or deforms under the applied loads.
• Low displacement areas (blue) indicate minimal movement or deformation.
• High displacement areas (red) suggest regions that experience significant deformation.
CHAPTER 6. SOFTWARES: ANALYSIS AND SIMULATIONS
62
Figure 6.13: nosecone stress
c. Strain
Equivalent Strain
It represents the overall deformation or elongation of the material under load.
• Low equivalent strain areas (blue) indicate minimal equivalent strain, approximately 4.61 × 10−16 .
• High equivalent strain areas (red) suggest potential failure points where the maximum equivalent strain
is around 2.8 × 10−4 . This value is still relatively low, suggesting that even the most strained parts of the
structure are experiencing deformation that is within acceptable limits and not indicative of imminent failure.
CHAPTER 6. SOFTWARES: ANALYSIS AND SIMULATIONS
63
Figure 6.14: nosecone strain
d. Safety factor
It is the ratio of the maximum load a structure can handle to the actual applied load. A safety factor greater than 1
means the structure is stronger than needed
Figure 6.15: Safety factor
CHAPTER 6. SOFTWARES: ANALYSIS AND SIMULATIONS
6.4.4
64
Simulation of the fins
The fin was constrained at its root chord and tabs (indicated in green) to simulate the reality of its real-world working
conditions. A parallel force, representing the air drag encountered during flight, to simulate aerodynamic forces
acting on the fin (indicated in purple).
Figure 6.16: the fins
6.4.5
Loads and Fixtures
Figure 6.17: Material proprieties
CHAPTER 6. SOFTWARES: ANALYSIS AND SIMULATIONS
Figure 6.18: Loads and Fixtures
6.4.6
Study Result
a. Stress
Figure 6.19: Fines stress
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CHAPTER 6. SOFTWARES: ANALYSIS AND SIMULATIONS
b. Displacement
Figure 6.20: Fines stress
c. Strain
Figure 6.21: Fins strain
d. Safety factor
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CHAPTER 6. SOFTWARES: ANALYSIS AND SIMULATIONS
67
Figure 6.22: Fins factor of safety
6.5
Computational fluid dynamics (CFD)
Computational fluid dynamics (CFD) is the numerical study of steady and unsteady fluid motion. The aerodynamic
performance of flight vehicles is of critical concern to airframe manufacturers, just as is the propulsive performance
of aircraft power plants, including those that are propeller-, gas turbine-, rocket, and electric driven. CFD is also
used to lessen the amount of physical testing that must be done to validate a design and measure its performance. It
is used to predict the drag, lift, noise, structural and thermal loads, combustion, etc., performance in aircraft systems
and subsystems.
It uses numerical methods and computer algorithms to model and simulate the behavior of fluids, such as liquids and
gases, to validate a design and measure its performance.
This approach involves first dividing space into finite elements or finite volumes, and then numerically solving the
essential equations of fluid dynamics, such as the Navier-Stokes equations, which describe fluid motion. (Computational Fluid Dynamics | Aerospace Engineering | UIUC.)
6.5.1
Grid Generation
Using ANSYS FLUENT Meshing, the domain is divided into elements and to ensure the accuracy and stability of the
solutions in critical zones (i.e. inlet and outlet regions, the boundary layer surroundings . . . etc), refining the mesh is
a necessity.
We constructed three distinct meshes with varying levels of refinement to evaluate the simulation accuracy.
Mesh 1: With 39140 elements.
Mesh 2: With 121468 elements.
Mesh 3: With 188056 elements.
CHAPTER 6. SOFTWARES: ANALYSIS AND SIMULATIONS
Figure 6.23: Mesh 1
Figure 6.24: Mesh 2
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CHAPTER 6. SOFTWARES: ANALYSIS AND SIMULATIONS
Figure 6.25: Mesh 3
69
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CHAPTER 6. SOFTWARES: ANALYSIS AND SIMULATIONS
6.5.2
Mesh Effect
We have successfully obtained the results for both drag force and drag coefficient using Fluent. To compare the
Mesh
Mesh 1
Mesh 2
Mesh 3
Drag Force (N)
307.81986
335.61980
415.57442
Cd
0.2285508
0.2498573
0.28221309
Table 6.1: Drag force and Cd for different meshes.
results, we computed the drag force using the drag coefficient for the body tube and nose cone from OpenRocket.
Fd =
1 2
ρv Cd A
2
• From Refprop, the air density ρ = 1.1348 kg/m3
• Vmax = 286 m/s
• Cd (nose cone + body tube) = 0.308, retrieved from OpenRocket
Figure 6.26: Drag coefficient from OpenRocket
D = 19.4 cm → A = 295.6 cm2 = 295.6 × 10−4 m2
Fd =
1
× 1.1348 × 2862 × 295.6 × 10−4 × 0.308 = 413.72 N
2
Figure 6.27: Drag force graph
CHAPTER 6. SOFTWARES: ANALYSIS AND SIMULATIONS
71
Figure 6.28: Drag Coefficient
From the graph we can say that mesh 3, featuring a refined structure and an optimal elements count, produces drag
coefficient and drag force statistics that closely match OpenRocket’s results.
Figure 6.29: Drag force from Ansys Fluent (Mesh 3)
Now that we’ve determined that Mesh 3 produces the best results, we will proceed with the simulation to further
analyze the flow phenomena using this mesh.
6.5.3
Solution settings and boundary conditions
In our ANSYS FLUENT simulation, we are primarily interested in investigating the behavior at the highest velocity
of 286 m/s (Mach 0.85). This velocity is the maximum that the rocket can reach during its trajectory and constitutes
a critical operating condition that we want to understand thoroughly.
Therefore, the k-ω SST model, in addition to the RANS equations, is used to solve the flow accurately.
The k-ω SST model is particularly effective for optimizing the aerodynamic performance of a rocket, due to its ability
to accurately resolve the complex phenomena that occur in the boundary layers, and to predict the flow separation
zones that cause drag-increasing turbulence.
• The k-ω SST model combines the advantages of the k-ω and k-ϵ models.
• The k-ω SST model is known to be more stable, which is advantageous for complex simulations of flows
around rockets.
6.5.4
Boundary conditions
The inlet boundary is defined with the normal speed (286m/s), and the outlet boundary is defined with the static
pressure.
6.5.5
Results
The obtained results are given by different contours of velocity, temperature and pressure
CHAPTER 6. SOFTWARES: ANALYSIS AND SIMULATIONS
72
a. Velocity
Figure 6.30: Velocity contour
Figure 6.31: Velocity at the tip of the nose cone
Comments
The velocity tends towards zero at the tip of the nose cone, where air stops momentarily. At this point a stagnation
point is created. This point since velocity is zero, thus creating a zone of maximum static pressure and static temperature.
At the outlet, a region of low or even zero velocity is observed. This is due to the formation of vortexes at the outlet,
resulting in local flow circulations. The increase from the lower values near the rocket’s surface to higher values in
the center of the flow is due to the air flow expansion which is accelerating due to the sudden change in crosssectional
area.
Velocity is uniform all along the flow zone; meanwhile the flow near the contact surfaces is decelerating which develops a thick boundary layer on the edges.
b. Static Temperature
CHAPTER 6. SOFTWARES: ANALYSIS AND SIMULATIONS
73
Figure 6.32: Static temperature contour
Comments
No noticeable heat transfer is seen and it’s normal the friction between air and the rocket does not yield an important
heating effect at these conditions (Fixed static temperature conditions were defined).
CHAPTER 6. SOFTWARES: ANALYSIS AND SIMULATIONS
74
c. Static Pressure
Figure 6.33: Static pressure contour
Comments
The pressure distribution around the rocket reveals distinct patterns, with a uniform pressure observed in the region,
suggesting a stable airflow and pressure distribution.
The stagnation point in front of a rocket’s cone is crucial point. This point since velocity is zero, thus creating a zone
of maximum static pressure.
Far from the exit section the separated flow that experienced vortexes meet again at the impact point where flow is
adjusted causing a high pressure.
6.5.6
Conclusion
The simulation results highlight the impact of the rocket’s shape on the flow characteristics. Further analysis could
focus on optimizing the rocket’s shape to minimize turbulence, reduce frictional losses, and promote stable flow.
The mesh produced for the simulation was carefully designed and validated through several steps. It shows excellent
compatibility with the results obtained from OpenRocket, which attests to its relevance for modeling the system.
Chapter 7
Manufacturing
There are several techniques available for this purpose, and they complement each other, allowing to switch from a
technique to another when the desired result is not met.
The choice of techniques highly depends on the materials used to construct the rocket and their manufacturability.
These techniques include:
a. 3D Printing.
b. Laser cutting, or mechanical cutting (Milling, band saw, cut-off saw...).
c. Lathe machining (Turning).
d. Wet Lay-up.
e. Casting.
f. Drill press (for necessary hole patterns).
For our fiberglass parts, a combination of these techniques was used to optimize performance and manufacturability.
Wet lay-up, in particular, was a primary method, along with other methods depending on the complexity of the part.
7.1
Manufacturing the Main Rocket
7.1.1
Manufacturing of the Body tube
The body tube of the rocket serves as its main structure and houses many of its essential components.
• Making a mandrel
To manufacture a fiberglass body tube, a mandrel is required. In our process, we used a PVC tube with an outer
diameter of 20 cm, a wall thickness of 4 cm, and a tolerance of 0.6 mm as the mandrel.
Given the length of the lower body tube (aft airframe) is 107 cm, the PVC tube was cut into two halves to facilitate
the wet lay-up process of the fiberglass (Fig. 7.1.). This step was essential to ensure easier handling and more precise
application of the fiberglass layers.
75
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CHAPTER 7. MANUFACTURING
Figure 7.1: The PVC pipe cutting
Figure 7.2: The half of the body tube
• Surface treatment
Before starting laying up the fiberglass, a surface treatment is needed, hence, the PVC tube was thoroughly cleaned
and coated with a release agent (Fig. 7.3.) to prevent the resin from sticking, making it easy to remove the finished
tube later, and for a smooth finish we applied Gel Coat.
Figure 7.3: Release agent
• Wet Lay-Up
Fiberglass cloths were cut for to the required dimensions, ensuring enough material to cover the entire surface of the
PVC tube.
After that, a mixture of resin, resin hardener, and hardening accelerator are all mixed together with portions according to the manufacturer’s instructions : the amount of resin specified, with 2% of that amount being hardener, and
2% of the resin amount being hardening accelerator.
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The fiberglass cloths were carefully laid over the mandrel halves, alternating between the resin mixture and fiberglass
layers. Each layer was saturated, starting with resin, followed by fiberglass, and finishing with another layer of resin,
ensuring no air bubbles or wrinkles were present using a bubbler remover roller.
• Curing
Once several layers are in place and desired thickness is achieved, the assembly is left to cure for quite some time so
the resin can harden and bond everything into one solid structure.
• Finishing
After the resin is cured the two portions of the lower body tube were carefully removed and assembled together using
resin, and then sealed with acrylic mastic sealant on the outside to avoid any gaps and ensure a clean finish.
The edges then are smoothed out, and sanded using sand paper so that any excess material or imperfections are
trimmed away.
Figure 7.4: Gel Coat application
Figure 7.5: Resin preparation
Figure 7.6: Resin application
Figure 7.7: Acrylic mastic sealant
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CHAPTER 7. MANUFACTURING
7.1.2
Nose cone
The nosecone which serves as a payload housing was 3D printed using PETG material with 50% infill. Due to the size
constraints of the available small 3D printers and the large size of the nose cone (97 cm), it was printed in sections
that were later assembled. (Fig. 7.9)
To ensure that the payload remains stationary (10x10x40) throughout the flight, we reduced the amount of space
inside the nosecone by creating a small housing for it that is already printed into the nose cone walls.
Figure 7.8: Nose cone part
7.1.3
Figure 7.9: Nose cone assembly
Fins
The process of manufacturing the same is almost identical to the body tube’s (Wet lay up method)
A smooth plate was cleaned and coated with the release agent and gel coat.
The mixture of the resin was prepared and the fiberglass clothes were cut into shape.
The wet lay-up process is then done alternating between fiberglass cloths and resin layers as previously done. (Fig.
7.10)
Again, once layers are in place and desired thickness is achieved, the assembly is then left to cure.
After everything had set, we sketched the fin shape into the plate and used an electric saw to cut accordingly. (Fig.
7.11)
To securely attach the fins to the body tube, we started by measuring and cutting a hole to the exact size needed. (Fig.
7.13)
Since there are three fins, we made sure the angle between them was 120 degrees within the tube using a fin jig we
made from hard cardboard. (Fig. 7.15)
A metallic "Shear Pin" was used to hold the fins in place from the inside, ensuring they were securely in place. (Fig.
7.16)
To further reinforce the attachment, we added more layers of fiberglass around the area.
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CHAPTER 7. MANUFACTURING
Figure 7.10: The piece of Fiberglass
Figure 7.11: Cutting the fins
Figure 7.12: Final result of the fins
Figure 7.13: Tang slotted airframe
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CHAPTER 7. MANUFACTURING
Figure 7.14: Fins attachment using shear pin
Figure 7.15: Fins attachment using Fiberglass
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CHAPTER 7. MANUFACTURING
7.1.4
Centering rings
A mold was made by forming a fiberglass model to the necessary measurements in order to construct the centering
rings for the motor. (Fig. 7.16)
Once the fiberglass model was complete, we used it to create our centering rings, applying the same materials and
process (the wet lay-up method).
The fiberglass was layered and cured, ensuring the rings were strong and precisely fitted to hold the motor securely
in place within the rocket. (Fig. 7.17)
Figure 7.16: The mold of centering rings
7.1.5
Figure 7.17: Final result of the centering rings
Bulkheads
Bulkheads serve multiple purposes, primarily protecting components from pressure caused by ejection charges and
providing attachment points for shock cords.
The thought process behind the fabrication of bulkheads was to go with the same idea used in making the fins.
The fiberglass plate of the desired thickness was carefully cut it into disks using a hole cutter (Fig. 7.18).
Then, the circularity of the disks was refined using a sharpening wheel to achieve a more precise shape.
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CHAPTER 7. MANUFACTURING
Figure 7.18: Hole cutter
7.1.6
Coupler
The coupler is used to assemble the upper body tube with the lower one, ensuring a proper connection between the
two parts. It was crafted using a PVC tube that measures 17cm in diameter that was cut to a desired length (28cm).
Bulkheads were then used to lock the PVC coupler openings, protecting the electronic components housed within.
These bulkheads do not only hold the coupler but also keep the black powder component firmly in place during
operation.
7.1.7
Parachute
For the parachutes, we began by cutting nylon cloth into the proper shape and sewing it into a hexagonal pattern.
(Fig. 7.19)
Holes were made for the shroud lines using a hole punch and metal eyelets, then we fastened the parachute shroud
lines when it was finished. (Fig. 7.21)
Both parachutes are stored in 3D-printed pistons designed to be propelled by the black powder charge. When ignited,
the charge pushes the pistons, ejecting the parachutes at the precise moment needed.
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CHAPTER 7. MANUFACTURING
Figure 7.19: Cutting the parachute parts
Figure 7.20: Parachute Sewing Process
Figure 7.21: Final Outcome of the Parachute Sewing Process
7.2
Manufacturing the Low altitude Rocket
All of the low altitude rocket components were mainly made from PVC, except for the nose cone which was 3D
printed.
7.2.1
Manufacturing of the Body tube
A 110 mm PVC tube was used as the rocket’s main body tube. The tube was cut to the exact length of 123 cm to meet
our design specifications. (Fig. 7.22)
7.2.2
Fins
To create the fins, we began with a PVC tube, which we heated to flatten. After flattening, we placed a heavy weight
on top to ensure a uniformly flat surface.
Once the tube was flattened, we sketched the fin shape and carefully cut it into the desired form using an electric saw.
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CHAPTER 7. MANUFACTURING
Figure 7.22: Leveling the body tube using a Level Bubble
To achieve a strong and secure fit to the body tube, the openings for the fins were cut into it and then firmly attached
using super glue. (Fig. 7.25)
Figure 7.23: Fins cutting
Figure 7.24: Tang slotted airframe
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CHAPTER 7. MANUFACTURING
Figure 7.25: Glowing the fins
7.2.3
Centering rings and bulkheads
Using the same PVC flattening procedure, we cut the centering rings and bulkheads using a hole cutter. (Fig. 7.26)
Centering rings and bulkheads were first attached to the inner motor tube then securely fastened to the body tube to
complete the assembly.
Figure 7.26: Centering rings
Figure 7.27: Bulkhead
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CHAPTER 7. MANUFACTURING
Figure 7.28: Glowing the fins
7.2.4
The Nose Cone
The nose cone was 3D printed using PLA material.
It was also printed in two sections, due to size constraints.
Figure 7.29: Nose Cone 3D printed
7.2.5
Parachute
The parachute was made the same way as the main rocket’s parachutes by sewing the cut pieces of cloth together. Once
completed, the metal eyelets and the shroud lines were attached to ensure proper functionality during deployment.
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CHAPTER 7. MANUFACTURING
7.3
Assembly
7.3.1
Body Tube Coupling
As mentioned before the airframe and the forward are connected with the coupler, and we make sure the coupler fits
within both body tubes.
Figure 7.30: The coupler
7.3.2
Parachute Packing and Placement
To prepare the parachutes, we folded them to prevent tangling during deployment.
We used a systematic folding technique to make sure they would unfold quickly and reliably. The shock cord was
tied to the suspension lines.
The folded parachutes were placed inside 3D-printed pistons, designed to hold them firmly in the body tube.
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CHAPTER 7. MANUFACTURING
7.3.3
Shock Cord Installation
We attached the shock cord to the coupler and the bulkhead, ensuring all sections of the rocket would stay connected
after deployment.
Figure 7.31: Shock cord installation
7.3.4
Electronics Integration
The electronic components were installed inside the coupler. We secured them to avoid any movement or disconnection caused by vibrations or acceleration during launch.
7.3.5
Black Powder Ejection System
For the ejection system, we placed the black powder inside glue sticks and then we put them inside pre-made holes
in the bulkheads on both sides of the coupler.
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CHAPTER 7. MANUFACTURING
Figure 7.32: Glue sticks
7.3.6
Final Assembly Check
We recheck all connections, and ensure that there’s no parts are loose or improperly assembled to confirm that the
rocket is fully prepared for deployment and recovery during flight.
For the law altitude rocket, we assembled the body tube with the fins securely attached, the 3D printed nose cone and
the inner motor tube.
Here is a photo of the final assembly for the two rockets:
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CHAPTER 7. MANUFACTURING
Figure 7.33: Assembled rocket
The painting process has been initiated, focusing on preparing the surface and applying the base layers. However,
the final detailing and finishing are still in progress. To achieve a polished and durable result, we will complete the
process using high-quality acrylic paint.
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CHAPTER 7. MANUFACTURING
Figure 7.34: Painting process
Chapter 8
Tests and Validations
8.1
Fiberglass Thermal Resistance
To confirm the rocket’s resistance, we tested the fiberglass, which is the primary material used in its fabrication, using
a blowtorch. This test used high temperatures to evaluate the fiberglass’s thermal resistance. The material resisted
the exposure without any significant deformation or failure, proving its capability to maintain performance during
flight.
Figure 8.1: Fiberglass Thermal Resistance Test
8.2
Ejection charge test
The ejection charge system was tested to ensure the efficiency of the black powder using the electronic circuit, and
that the ejection mechanism would activate correctly and to determine the necessary amount of powder.
Although the test was simple, it was effective in proving the system’s durability and performance, guaranteeing that
the ejection charge worked well during the rocket’s flight. A full-scale ejection test of the entire rocket will be conducted soon to further validate the system’s performance.
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CHAPTER 8. TESTS AND VALIDATIONS
93
Figure 8.2: Ejection charge test
8.3
Deployment test
In order to closely monitor the parachute’s deployment, we tested it by launching it from a significant height. Our
primary objective was to ensure that it opened correctly and descended in a controlled manner, and also, we want to
confirm the parachute’s behavior as it unfurled and filled with air.
8.4
Tensile tests
Tensile testing on various materials revealed their elastic and plastic properties, as well as their resistance to rupture.
In order to validate the properties utilized in simulations for structural calculations, these evaluations are essential.
Tensile tests are applied to the following materials:
• PLA
• PETG
• Fiberglass
• Shock Cord
• Shroud Lines
CHAPTER 8. TESTS AND VALIDATIONS
94
Figure 8.3: Tensile testing machine
8.4.1
PLA & PETG
For both a tensile test was carried out on PLA and PETG with an infill of 20%. The results show that :
• PLA with both 200% infill can withstand up to 3.14kN.
• PETG with both 20% infill can withstand up to 4.21kN.
The results indicate that both materials have good strength. Our observations revealed that a 20% infill makes the
structure very light, while going up to 80% makes it quite heavy. To strike a balance between weight and strength, we
decided on an intermediate solution with 50% infill.
Figure 8.4: PLA & PETG test specimens
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CHAPTER 8. TESTS AND VALIDATIONS
Figure 8.5: Force variation over time for PLA
8.4.2
Figure 8.6: Force variation over time for PETG
Fiberglass composite
Fiberglass as a composite material, it is important to understand its characteristics and resistance, so we made a sheet
of fiberglass in 5 layers with resin, then cut samples to test them.
We determined that the fiberglass composite can withstand up to 5.5 kN.
Please note that we couldn’t provide a picture or a graph for the results due to a technical problem in the tensile
machine, which led to the loss of the data.
Figure 8.7: Fiberglass composite test specimens
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CHAPTER 8. TESTS AND VALIDATIONS
8.4.3
Shock cord and shroud lines
After thorough testing, we determined that the shock cord can withstand an impressive load of up to 5.5 kN, whereas
the shroud lines resist up to 7kN and since this is more than enough, the test was stopped before the rupture point
because it took quite a bit of time to complete.
Overall, the tensile tests were successful, confirming that the shock cord and the shroud lines are capable of resisting
the applied forces and maintaining their structural integrity.
Figure 8.8: Shroud line test
Conclusion
Project Self-Evaluation
This project has been an enriching experience for our team, which has given us priceless knowledge on time management, problem-solving, and project management.
We were given the chance to put our academic knowledge to use in a real-world setting, which strengthened our
comprehension and allowed us to explore new concepts and techniques.
Even with our achievements, every project comes with its challenges. One of the main difficulties we faced was related to delays in material procurement so we had to make adjustments without changing the program as a whole
because the ordering procedure took longer than expected, which affected our timeline.
Also, the fabrication trials and the process of confirming materials and methods demanded attention, which added
to the project’s complexity.
Another major obstacle was time limits, which forced us to operate under pressure to fulfill deadlines.
One of the most valuable aspects of this journey was the way we grew as a team. . We discovered how to work well
together, utilizing one another’s skills and strengths while helping one another overcome obstacles.
Through this experience, we learned how to improve our communication, assign duties effectively, and adjust as a
team to solve challenges.
Conclusion
In conclusion, the journey of designing and building a high-power rocket for CIAA has been both challenging and
rewarding. It has gained us a wide knowledge on various topics, from the historical context of rocketry, to the
theoretical considerations and structural analysis involved in our design process.
Our design choices were made with careful consideration, utilizing various analytical methods and rule of thumbs,
as well as powerful tools such as OpenRocket, Rocksim, SOLIDWORKS, and ANSYS FLUENT.
These choices were validated through simulations and structural strength calculations, which not only confirmed the
robustness of our design but also provided us with valuable insights and learning experiences.
This project has been more than just about building a rocket; it has been about applying theoretical knowledge in a
practical context, learning from both successes and failures, and continuously striving for improvement.
97
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22, 2024, from https://aerospace.illinois.edu/research/research-areas/computational-fluid-dynamics
[3] Howard, Z. (2011, July 19). How To Calculate Fin Flutter Speed. Peak of Flight, (291), 3.
https://www.apogeerockets.com/education/downloads/Newsletter291.pdf
[4] Gasmire, C. Model Rocket Fins 101: Purpose, Shape, Size, and Placement – The Model Rocket. The Model
Rocket. Retrieved May 18, 2024, from https://themodelrocket.com/model-rocket-fins-101-purpose-shapesize-and-placement/
[5] Manfredo,
John.
“Properly
Sizing
Parachutes
for
Your
Rockets.”
Apogee
https://apogeerockets.com/education/downloads/Newsletter149.pdf. Accessed 21 November 2024.
Rockets,
[6] Model Rocket Guides. AustinRockets. Retrieved May 18, 2024, from https://austinrockets.org/how-to-findoptimal-fin-size-for-model-rockets/
[7] Rockets 101 | National Geographic. (2019, July 12). Youtube. Retrieved May 18, 2024, from
https://youtu.be/1yBwWLunlOM?si=1InIK-HbCmn1gCT3&t=46
[8] Propulsion System. (2023, November 20). Glenn Research Center | NASA. Retrieved May 18, 2024, from
https://www1.grc.nasa.gov/beginners-guide-toaeronautics/propulsion-system/
[9] Richard Nakka’s Experimental Rocketry Site. (2022, July 20). Richard Nakka’s Experimental Rocketry Site. Retrieved May 21, 2024, from https://www.nakkarocketry.net/RD_body.html
[10] Sizing Fins | STAR Public. (2020, November 3). Space Technologies and Rocketry. Retrieved May 18, 2024, from
https://rocketry.gitbook.io/public/tutorials/airframe/sizing-fins
[11] The University of Maine / School of Marine Sciences Physical Oceanography Group. Variable Description. Retrieved May 20, 2024, from http://gyre.umeoce.maine.edu/data/gomoos/buoy/php/variable_description.php
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[12] Rick Hanke, Maui Ultra Fins - Fin Secrets, 2011, https://www.mauiultrafins.shop/mediafiles/Bilder/secrets/Maui_Ultra_Fins_Fin_
04-09.pdf
[13] Sensor Embedded, Arduino Development Boards,
development-boards.html
98
2024,
https://www.sensorembedded.in/arduino-
APPENDIX A: Drawing of the design (
Low altitude rocket )
99
BIBLIOGRAPHY
100
BIBLIOGRAPHY
101
BIBLIOGRAPHY
102
BIBLIOGRAPHY
103
APPENDIX B: Virtual maquette of the
main rocket
104
BIBLIOGRAPHY
105
BIBLIOGRAPHY
106
APPENDIX C: Barrowman method for
calculating center of pressure
107
APPENDIX D: Static Stability Criteria
Static Stability Criteria
To ensure optimal rocket stability, the following criteria must be met:
• Static Margin (MS)
The static margin should be within the range:
2 < M S = 2.2 < 6
• Aerodynamic Fineness (L/D)
The length-to-diameter ratio of the rocket should be between:
10 < L/D = 12.63 < 35
• Ground-Level Velocity
The velocity at ground level should be less than 10 m/s:
Vg = 8.46 m/s < 10 m/s
• Ramp Exit Velocity
The minimum velocity of the rocket at the launch rail exit should exceed 20 m/s:
Ve = 25.2 m/s > 20 m/s
108
APPENDIX E: Rocket Materials List
Table 8.1: Rocket Materials List: Mechanical and Electronic Components
Component
Quantity
Mechanical Components
3D Printer
/
Cutter
30m
Nylon Fabric
15m
Fiberglass
2 Kilo
Balance
/
Epoxy Resin
/
Polyester Resin
8 Kilo
Adhesive Tape
/
Scissors
/
Sandpaper
/
Paintbrushes
/
Solvent (Acetone)
1L
Release Agent
5L
Mixing Sticks
/
Gloves
/
Shock Cord
10m
Welding Pen
/
Polisher
/
Drill
/
PVC Tube 20cm
3m
PVC Tube 11cm
2m
Spray Paint (Acrylic)
/
Dremel Tool
/
Hole Saw
160g
Black Powder
/
Gel Coat
/
SuperGlue
/
Ruler
/
Electric Saw
/
Sharpening Wheel
/
Mastic
/
Hardener
/
Hardening Accelera- /
tor
Calipers
/
Price per Unit
(DZD)
Total
(DZD)
/
100–550
400
800
/
250–280
450–800
350
150–600
75
250–450
2000
1100
250
150
250
/
/
/
320–380
320–380
350
/
19
/
/
35
/
/
/
/
/
/
/
Price
Mass per Unit
(g)
Total Mass (g)
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
/
Continued on next page
109
110
BIBLIOGRAPHY
Table 8.1: Rocket Materials List (continued)
Component
Quantity
Price per Unit
(DZD)
Total
(DZD)
Miscellaneous Hardware (screw eyelets,
shear pins, threaded
inserts, quick links,
eye bolts, and removable rivets)
/
/
Electronic Components
Arduino Mega
1
Module GPS NEO8M 1
Bosch BMP BMP280 1
LoRa 32 V2
2
SD Card 2GB
1
Module SPI pour 1
Carte
Mémoire
Micro SD TF
3.7V
Rechargeable 5
Lithium-ion Battery
Convertisseur abais- 1
seur LM2596 SMD
3A
MPU6050
1
Total (Electronics)
Mass per Unit
(g)
Total Mass (g)
/
/
/
3700
3500
1000
5000
650
350
3700
3500
1000
10000
650
350
37
19
3
100
5
10
37
19
3
200
5
10
1000
5000
50
250
600
600
20
20
800
800
3
3
23600
Price