ISTANBUL TECHNICAL UNIVERSITY ⋆ GRADUATE SCHOOL

PRELIMINARY DESIGN TOOL FOR LAUNCH VEHICLES
WITH LIQUID ROCKET ENGINES

M.Sc. THESIS

Tegin Berkay BUDAK

Department of Defense Technologies

Defense Technologies Programme

JUNE 2025





ISTANBUL TECHNICAL UNIVERSITY ⋆ GRADUATE SCHOOL

PRELIMINARY DESIGN TOOL FOR LAUNCH VEHICLES
WITH LIQUID ROCKET ENGINES

M.Sc. THESIS

Tegin Berkay BUDAK
(514201036)

Department of Defense Technologies

Defense Technologies Programme

Thesis Advisor: Prof. Dr. Alim Rüstem ASLAN

JUNE 2025





İSTANBUL TEKNİK ÜNİVERSİTESİ ⋆ LİSANSÜSTÜ EĞİTİM ENSTİTÜSÜ

SIVI ROKET MOTORLU FIRLATMA ARAÇLARI İÇİN
ÖN TASARIM ARACI

YÜKSEK LİSANS TEZİ

Tegin Berkay BUDAK
(514201036)

Savunma Teknolojileri Departmanı

Savunma Teknolojileri Programı

Tez Danışmanı: Prof. Dr. Alim Rüstem ASLAN

HAZİRAN 2025





Tegin Berkay BUDAK, a M.Sc. student of ITU Graduate School student ID 514201036
successfully defended the thesis entitled “PRELIMINARY DESIGN TOOL FOR LAUNCH
VEHICLES WITH LIQUID ROCKET ENGINES”, which he/she prepared after ful-
filling the requirements specified in the associated legislations, before the jury whose
signatures are below.

Thesis Advisor : Prof. Dr. Alim Rüstem ASLAN ..............................
Istanbul Technical University

Jury Members : Assist. Prof. Dr. Cuma YARIM ..............................
Istanbul Technical University

Assist. Prof. Dr. Erdem Onur ÖZYURT ..............................
Turkish-German University

Date of Submission : 25 May 2025
Date of Defense : 17 June 2025

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vi



To my only and everlasting love,

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viii



FOREWORD

Firstly, I would like to express my deepest gratitude to my thesis advisor, Prof.
Dr. Alim Rüstem ASLAN, for his invaluable guidance, insightful feedback, and
unwavering support throughout the course of this research. His expertise and
encouragement have been instrumental in shaping both the direction and the quality
of this work.

I am also grateful to my friends and family for their moral support and understanding.
In particular, I wish to thank my parents for fostering my curiosity and for always
believing in my potential.

Lastly, I would like to express my sincere gratitude to my wife, Betül YILDIRIM
BUDAK, whose love, support, and patience got me through every difficult moment.
My greatest strength has come from her unshakable confidence in me.

June 2025 Tegin Berkay BUDAK
(Astronautical Engineer)

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TABLE OF CONTENTS

Page

FOREWORD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ix
TABLE OF CONTENTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xiii
ABBREVIATIONS. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xv
SYMBOLS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xvii
LIST OF TABLES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xxi
LIST OF FIGURES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xxv
SUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .xxvii
ÖZET . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xxix
1. INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

1.1 Purpose of Thesis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1
1.2 Similar Works in Literature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
1.3 Hypothesis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6
1.4 Thesis Outline . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

2. FUNDAMENTALS AND EVOLUTION OF LAUNCH VEHICLE
DESIGN . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

2.1 Launch Vehicles: Historical Perspective and Current Trends. . . . . . . . . . . . . . . 9
2.2 Launch Vehicle Fundamentals . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20

2.2.1 Launch vehicle classification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20
2.2.2 Components of launch vehicles . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22
2.2.3 Launch vehicle mass elements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27
2.2.4 Launch vehicle performance. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28

2.2.4.1 Delta-V requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28
2.2.4.2 Specific impulse (Isp) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29
2.2.4.3 Payload fraction (π) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32
2.2.4.4 Structural factor (σ ) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33

2.2.5 Staging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36
2.2.6 Launch sites . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38

2.3 Rocket Engine Fundamentals . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40
2.3.1 Liquid rocket engines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 46

2.3.1.1 Liquid rocket engine propellants . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 49
2.3.1.2 Cycles of liquid rocket engines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 49

2.4 Launch Vehicle Design Methodology. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 59
2.4.1 Stages of design and development . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 60
2.4.2 Preliminary design phase . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62

3. TOOL DEVELOPMENT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65
3.1 Tool Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65
3.2 Mission Energy Requirements Sub-model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 66

3.2.1 Operational constraints of the model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 66
3.2.2 Model steps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 68

3.3 Propulsion Sub-model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 73
3.4 Stage Delta-V (∆V ) Distribution Sub-model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 78

3.4.1 Stage optimum ∆V optimization . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 81
3.4.2 Stage mass distributions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 83
3.4.3 Thrust-to-weight ratio and maximum acceleration constraints . . . . . . . . 86

3.5 Mass Properties Sub-model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 88
3.5.1 Engine properties calculation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 89
3.5.2 Propellant mass and volume calculation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 95
3.5.3 Propellant tank concepts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 98
3.5.4 Length properties calculation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 100
3.5.5 Diameter scan range suggestions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 103

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3.5.6 Mass properties calculation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 106
3.5.6.1 Propellant tanks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 106
3.5.6.2 Insulation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 109
3.5.6.3 Interstage, intertank, forward skirt and aft skirt . . . . . . . . . . . . . . . . . . . . 111
3.5.6.4 Liquid rocket engine . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 114
3.5.6.5 Thrust frame . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 123
3.5.6.6 Thrust vector control (TVC) system. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 125
3.5.6.7 Pressurization and feed system . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 126
3.5.6.8 Avionics, primary power and wiring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 127
3.5.6.9 Payload fairing. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 129

3.6 Cost Properties Sub-model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 137
3.6.1 Development cost estimation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 137

3.6.1.1 Cost scaling factors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 138
3.6.1.2 System engineering and integration factor ( f0) . . . . . . . . . . . . . . . . . . . . 138
3.6.1.3 Technical quality factor ( f2) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 139
3.6.1.4 Team experience factor ( f3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 140
3.6.1.5 Parallel organization factor ( f7). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 140
3.6.1.6 Development schedule factor ( f6) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 141
3.6.1.7 Development type coefficient ( fxx,stage or fxx,eng) . . . . . . . . . . . . . . . . . . 142

3.6.2 Production cost estimation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 142
3.6.2.1 Engine production cost . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 143
3.6.2.2 Cost reduction factor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 144

4. TOOL VALIDATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 147
4.1 Mission Energy Requirements Validation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 148
4.2 Propulsion Validation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 150
4.3 Mass Properties Validation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 153

4.3.1 Data used for mass element validation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 153
4.3.2 Propellant tanks and insulation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 154
4.3.3 Interstage, intertank, forward skirt and aft skirt . . . . . . . . . . . . . . . . . . . . . . . . 156
4.3.4 Liquid rocket engine and gimbal (TVC) system . . . . . . . . . . . . . . . . . . . . . . . 157
4.3.5 Thrust frame . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 160
4.3.6 Pressurization and feed system . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 160
4.3.7 Avionics, primary power and wiring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 163
4.3.8 Payload fairing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 164
4.3.9 Reserve, start and unused propellant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 167

4.4 Cost Properties Validation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 168
5. TOOL APPLICATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 171

5.1 Case Study 1: Design Strategy for Small Launch Vehicle to Deliver
10,000 kg Total Payload . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 171

5.1.1 1000 kg payload vehicle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 175
5.1.2 500 kg payload vehicle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 179
5.1.3 Comparison . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 185
5.1.4 Conclusion. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 189

5.2 Case Study 2: Sensitivity Analysis of First Stage T/W Ratio and Engine
Clustering in Multistage Launch Vehicle Design. . . . . . . . . . . . . . . . . . . . . . . . . . . . 190

5.2.1 System mass margin variation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 191
5.2.2 Thrust frame mass variation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 192
5.2.3 Engine mass variation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 192
5.2.4 Thrust vector system mass variation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 193
5.2.5 Burn time vs. T/W ratio. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 194
5.2.6 Conclusion. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 195

5.3 Case Study 3: Super-Heavy Launch Vehicle of Türkiye . . . . . . . . . . . . . . . . . . . . 199
5.3.1 Launch site comparison . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 199
5.3.2 Engine candidates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207
5.3.3 Launch vehicle concept candidates. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208
5.3.4 Configuration 1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212
5.3.5 Configuration 2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 214
5.3.6 Configuration 3 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 215
5.3.7 Configuration 4 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 216
5.3.8 Selected configurations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 218

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5.3.9 Cost estimation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 219
6. CONCLUSIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 223

6.1 Further Studies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 226
REFERENCES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 229
APPENDICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 237

APPENDIX A : Packing of circles in a larger circle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 239
APPENDIX B : Mission scenarios based on user input . . . . . . . . . . . . . . . . . . . . . . . . . . . 241
APPENDIX C : Orbital calculations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 245

CURRICULUM VITAE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 249

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ABBREVIATIONS

AIAA : American Institute of Aeronautics and Astronautics
CEA : Chemical Equilibrium with Applications
ECO : Engine Cut-Off
EX : Expander Cycle
GG : Gas Generator Cycle
GEO : Geo-Synchronous Earth Orbit
GLOM : Gross Lift-Off Mass
GLOW : Gross Lift-Off Weight
GTO : Geostationary Transfer Orbit
IMU : Inertial Measurement Unit
ICBM : Intercontinental Ballistic Missile
ISS : International Space Station
JP : Japan
KR : South Korea
LCC : Life Cycle Cost
LCH4 : Liquid Methane
LEO : Low Earth Orbit
LEO : Low-Earth Orbit
LH2 : Liquid Hydrogen
LOX : Liquid Oxygen
LRE : Liquid Rocket Engine
MER : Mass Estimation Relationship
MGA : Mass Growth Allowance
MDO : Multidisciplinary Design Optimization
MEO : Middle-Earth Orbit
MLI : Multi-Layer Insulation
MPa : Megapascal
NASA : National Aeronautics and Space Administration
N2O4 : Dinitrogen Tetroxide
O/F : Oxidizer-to-fuel Ratio
PEO : Polar Earth Orbit
PLA : Payload Adapter
PMD : Propellant Management Devices
PMF : Propellant Mass Fraction
RSE : Residual Standard Error
RP-1 : Rocket Propellant – 1
SC : Staged Combustion Cycle
SL : Sea Level

xv



SSME : Space Shuttle Main Engine
SSO : Sun-Synchronous Earth Orbit
TR : Türkiye
T/W : Thrust-to-Weight Ratio
TVC : Thrust Vector Control
UDMH : Unsymmetrical Dimethylhydrazine
US : United States
V-2 : Vergeltungswaffe 2, Eng. Retribution Weapon 2

xvi



SYMBOLS

∆V : Delta-V, Velocity Change

xvii



xviii



LIST OF TABLES

Page

Table 2.1 : V-2 missile specifications. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12
Table 2.2 : V-2 missile liquid engine specifications. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13
Table 2.3 : Launch vehicle classification by payload capacity. . . . . . . . . . . . . . . . . . . . . 21
Table 2.4 : Launch vehicle classification by mission types and their applications. 21
Table 2.5 : Launch vehicle components (1 to 7) continued. . . . . . . . . . . . . . . . . . . . . . . . . 24
Table 2.6 : Launch vehicle components (8 to 14) continued. . . . . . . . . . . . . . . . . . . . . . . 24
Table 2.7 : Launch vehicle components (15 to 21) continued. . . . . . . . . . . . . . . . . . . . . . 25
Table 2.8 : Launch vehicle components (22 to 28) continued. . . . . . . . . . . . . . . . . . . . . . 26
Table 2.9 : Typical Delta-V requirements for various space missions. . . . . . . . . . . . . 29
Table 2.10 :Typical specific impulse values for different propulsion systems. . . . . . 31
Table 2.11 :Payload fraction of various launch vehicles. . . . . . . . . . . . . . . . . . . . . . . . . . . . 32
Table 2.12 :Structural factor (σ ) of first stages of various launch vehicles. . . . . . . . . 34
Table 2.13 :Structural factor (σ ) of second stages of various launch vehicles. . . . . 35
Table 2.14 :Structural factor (σ ) of third stages of various launch vehicles. . . . . . . . 35
Table 2.15 :Global launch sites and their orbital capabilities. . . . . . . . . . . . . . . . . . . . . . . 39
Table 2.16 :Comparison of chemical propulsion systems. . . . . . . . . . . . . . . . . . . . . . . . . . . 41
Table 2.17 :Typical characteristics of LREs. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 47
Table 2.18 :Typical characteristics of LREs continued. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 48
Table 2.19 :AIAA recommended mass growth allowance (%) . . . . . . . . . . . . . . . . . . . . . 62
Table 3.1 : User inputs for the propulsion sub-model. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 74
Table 3.2 : Example input for the propulsion sub-model. . . . . . . . . . . . . . . . . . . . . . . . . . . 76
Table 3.3 : Example output of the propulsion sub-model. . . . . . . . . . . . . . . . . . . . . . . . . . . 77
Table 3.4 : Sample input of launch vehicle stage ∆V distribution sub-model. . . . . 79
Table 3.5 : Pseudo-code for the bisection method. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 83
Table 3.6 : Pseudocode for mass calculation in stage distribution. . . . . . . . . . . . . . . . . 86
Table 3.7 : Sample output of launch vehicle stage ∆V distribution sub-model. . . . 88
Table 3.8 : Chamber characteristic lengths (L∗) for different propellant

combinations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 92
Table 3.9 : Method 1 rocket engine fit curve coefficients. . . . . . . . . . . . . . . . . . . . . . . . . . . 116
Table 3.10 :Liquid rocket engine data from various references. . . . . . . . . . . . . . . . . . . . . 119
Table 3.11 :Cost scaling factors for development and production estimation. . . . . . 138
Table 3.12 :Team experience factor for cost estimation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 140
Table 3.13 :Development type coefficient for cost estimation. . . . . . . . . . . . . . . . . . . . . . 142
Table 4.1 : Design velocities (in m/s) by launch vehicle.. . . . . . . . . . . . . . . . . . . . . . . . . . . 148
Table 4.2 : Comparison of estimated ∆V with different correction methods. . . . . . 149
Table 4.3 : Input parameters for propulsion sub-model validation. . . . . . . . . . . . . . . . . 151
Table 4.4 : Comparison of CEA and sub-model outputs for case 1. . . . . . . . . . . . . . . . 152
Table 4.5 : Comparison of CEA and sub-model outputs for case 2. . . . . . . . . . . . . . . . 152

xix



Table 4.6 : Estimated mass values for Saturn V tanks. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 155
Table 4.7 : Errors in mass estimation for Saturn V tanks (%). . . . . . . . . . . . . . . . . . . . . . 155
Table 4.8 : Estimated mass values for Ariane 5 oxidizer and fuel tanks. . . . . . . . . . . 155
Table 4.9 : Estimation errors for Ariane 5 oxidizer and fuel tanks. . . . . . . . . . . . . . . . . 156
Table 4.10 :Estimated mass values for structural components. . . . . . . . . . . . . . . . . . . . . . 157
Table 4.11 :Estimation errors for structural components. . . . . . . . . . . . . . . . . . . . . . . . . . . . 157
Table 4.12 :Estimated mass values of Ariane 5 for engines with TVC. . . . . . . . . . . . . 158
Table 4.13 :Estimation errors of Ariane 5 for engines with TVC. . . . . . . . . . . . . . . . . . . 158
Table 4.14 :Estimated mass values for Saturn V stage engines. . . . . . . . . . . . . . . . . . . . . 158
Table 4.15 :Estimation errors for Saturn V stage engines. . . . . . . . . . . . . . . . . . . . . . . . . . . 159
Table 4.16 :TVC mass estimation for Saturn V stages. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 160
Table 4.17 :Pressurization system mass for various launch vehicle stages. . . . . . . . . 161
Table 4.18 :Validation of propellant system mass estimation. . . . . . . . . . . . . . . . . . . . . . . 163
Table 4.19 :Estimated avionics mass values and errors for Saturn V and Ariane 5. 164
Table 4.20 :Launch vehicles and their respective payload fairing masses. . . . . . . . . . 165
Table 4.21 :Unused reserve and propellant estimation for Saturn V and Ariane 5. 167
Table 4.22 :Development costs for selected launch vehicles (FY 2006 $). . . . . . . . . 168
Table 4.23 :Validation of estimated development costs versus actual costs. . . . . . . . 169
Table 5.1 : Ideal engine performance parameters used in the study. . . . . . . . . . . . . . . . 173
Table 5.2 : Specifications of the two-stage launch vehicle for 1000 kg payload. . 179
Table 5.3 : Maximum stage 2 PMF to maintain ≥ 15% structural mass margin. . 182
Table 5.4 : Minimum stage 2 PMF values to achieve payload fraction ≥ 1%. . . . . 183
Table 5.5 : Acceptable stage 2 PMF ranges for margin and payload fraction

compliance. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 184
Table 5.6 : Specifications of the two-stage launch vehicle for 500 kg payload. . . . 184
Table 5.7 : Comparison of two-stage launch vehicles for 500 kg and 1000 kg

payloads. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 185
Table 5.8 : Assumptions for cost factor ( f ) analysis of launch vehicle elements. . 186
Table 5.9 : Assumptions for production numbers of stages and engines. . . . . . . . . . . 187
Table 5.10 :Assumed p-factors for cost reduction analysis. . . . . . . . . . . . . . . . . . . . . . . . . 187
Table 5.11 :Corresponding f4 values for cost reduction analysis. . . . . . . . . . . . . . . . . . . 188
Table 5.12 :Total development and production costs of the launch vehicles. . . . . . . 188
Table 5.13 :Launch vehicle parameters used in the case study. . . . . . . . . . . . . . . . . . . . . . 191
Table 5.14 :Geographic coordinates of candidate launch sites. . . . . . . . . . . . . . . . . . . . . . 200
Table 5.15 :Selected launch azimuths, corresponding parking orbit inclinations,

and total mission ∆V requirements for each candidate launch site. . . . 204
Table 5.16 :Candidate engine configurations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208
Table 5.17 :Ideal performance specifications of the six reference rocket engines.. 208
Table 5.18 :Specifications of launch vehicle configurations. . . . . . . . . . . . . . . . . . . . . . . . 211
Table 5.19 :Mass breakdowns of launch vehicle configurations. . . . . . . . . . . . . . . . . . . . 211
Table 5.20 :Engine counts and mass margins for selected configurations 1 and 4. 218
Table 5.21 :Stage thrust parameters for selected configurations 1 and 4. . . . . . . . . . . 219
Table 5.22 :Assumptions for cost factor ( f ) analysis of launch vehicle elements. . 219
Table 5.23 :Production quantities by configuration and stage. . . . . . . . . . . . . . . . . . . . . . . 220
Table 5.24 :Development and production costs for selected configurations (in

MYR). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 220

xx



Table 5.25 :Development and production costs converted to USD (in billions),
using 1 MYR = 200 000 USD. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 220

Table A.1 : Densest packing possible minimum diameter.. . . . . . . . . . . . . . . . . . . . . . . . . . 239

xxi



xxii



LIST OF FIGURES

Page

Figure 2.1 : Tsiolkovsky’s work, 1914 edition. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10
Figure 2.2 : Left. Goddard’s liquid propellant rocket. Right. Goddard’s

rocket’s components. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11
Figure 2.3 : V-2 ballistic missile. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12
Figure 2.4 : V-2 engine. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13
Figure 2.5 : R-7 Semyorka to Soyuz launch vehicle. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14
Figure 2.6 : Side and bottom views of N-1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15
Figure 2.7 : Family of Saturn rockets.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16
Figure 2.8 : Saturn V. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16
Figure 2.9 : Space Shuttle launch. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17
Figure 2.10 :Buran orbiter. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17
Figure 2.11 :Launch costs. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18
Figure 2.12 :Launch vehicle components. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23
Figure 2.13 :Comparison of Delta-V requirements for different space missions. . 30
Figure 2.14 :Structural factor σ as a function of propellant mass for various

launch vehicle stages. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36
Figure 2.15 :Launch vehicle staging. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 37
Figure 2.16 :Rocket engine control volume.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42
Figure 2.17 : Isp with the changing O/F. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43
Figure 2.18 :Tc of LOX-LH2 with the changing O/F. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 44
Figure 2.19 :Tc of LOX-RP1 with the changing O/F. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45
Figure 2.20 :Tc of N2O4-MMH with the changing O/F. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45
Figure 2.21 :Pressure fed cycle. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50
Figure 2.22 :OMS engine outer view.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51
Figure 2.23 :OMS engine schematic of feed lines. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51
Figure 2.24 :Gas generator cycle. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52
Figure 2.25 :F-1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53
Figure 2.26 :Merlin-1D. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53
Figure 2.27 :Typical (fuel-rich) staged combustion cycle. . . . . . . . . . . . . . . . . . . . . . . . . . . 55
Figure 2.28 :Full flow staged combustion cycle. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55
Figure 2.29 :Open expander cycle. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 57
Figure 2.30 :Closed expander cycle. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 57
Figure 2.31 :Electric pump cycle. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 58
Figure 2.32 :9 x Rutherford engines seen on the first stage of Electron LV. . . . . . . . 58
Figure 2.33 :Launch vehicle design process flowchart. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 59
Figure 2.34 :Design process/life-cycle flow.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 60
Figure 2.35 :Mass growth trends across various launch vehicle programs. . . . . . . . . 61

xxiii



Figure 2.36 :Breakdown of launch vehicle wet mass.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 63
Figure 3.1 : Preliminary design tool flowchart. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65
Figure 3.2 : Mission energy requirements sub-model. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 68
Figure 3.3 : Launch azimuth with corresponding inclination. . . . . . . . . . . . . . . . . . . . . . 69
Figure 3.4 : Mission scenario decision and calculation.. . . . . . . . . . . . . . . . . . . . . . . . . . . . 70
Figure 3.5 : Mission ∆V summary. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 72
Figure 3.6 : Propulsion sub-model flow chart. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 73
Figure 3.7 : Change of Isp with different O/F for LOX – RP1. . . . . . . . . . . . . . . . . . . . 75
Figure 3.8 : Change of Isp with different O/F for LOX – LH2. . . . . . . . . . . . . . . . . . . . 76
Figure 3.9 : Stage Delta-V (∆V ) distribution sub-model steps. . . . . . . . . . . . . . . . . . . . . 78
Figure 3.10 :Mass properties sub-model steps. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 89
Figure 3.11 :Contraction ratio vs throat diameter.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 93
Figure 3.12 :Dimensions of liquid rocket engine. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 96
Figure 3.13 :Start-up transient of a liquid rocket engine. . . . . . . . . . . . . . . . . . . . . . . . . . . . 97
Figure 3.14 :Propellant tank concept. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 99
Figure 3.15 :Aspect ratio of propellant tank dome. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 99
Figure 3.16 :Dimensions of launch vehicle. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 102
Figure 3.17 :Diameters of launch vehicles. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 104
Figure 3.18 :L/D of launch vehicles. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 106
Figure 3.19 :Tank mass correlation with volume. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 108
Figure 3.20 : Insulation areal density based on propellant storage temperature. . . . 110
Figure 3.21 :Mass of structures. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 114
Figure 3.22 :Method 2 of mass estimation for liquid rocket engines. . . . . . . . . . . . . . . 117
Figure 3.23 :Expander cycle regression curve. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 119
Figure 3.24 :Staged combustion cycle regression curve. . . . . . . . . . . . . . . . . . . . . . . . . . . . 120
Figure 3.25 :Gas generator cycle regression curve for Fvac,kN < 400kN. . . . . . . . . . . 121
Figure 3.26 :Gas generator cycle regression curve for Fvac,kN > 400kN. . . . . . . . . . . 121
Figure 3.27 :Altas V 401 mission profile. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 130
Figure 3.28 :Delta IV M+(5,4) mission profile. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 131
Figure 3.29 :Delta IV M+(5,4) mission profile. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 132
Figure 3.30 :Dimensions of the payload fairing calculation. . . . . . . . . . . . . . . . . . . . . . . . 135
Figure 3.31 :Observed h1/h2 ratios from various launch vehicle fairing designs. . 136
Figure 3.32 :Reference net mass fraction for cryo stages. . . . . . . . . . . . . . . . . . . . . . . . . . . 139
Figure 3.33 :Reference net mass fraction for storable or semi-cryo stages. . . . . . . . 140
Figure 3.34 :Development schedule factor. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 141
Figure 3.35 :Learning factor. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 145
Figure 3.36 :Production cost reduction factor. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 145
Figure 4.1 : Thrust frame mass estimation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 161
Figure 4.2 : Pressurization system mass estimation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 162
Figure 4.3 : Payload fairing mass and estimation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 165
Figure 4.4 : Payload fairing volume estimation from payload mass capability. . . . 166
Figure 5.1 : Payload fraction for various PMF values. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 174
Figure 5.2 : 1000 kg PL vehicle PMF values vs stage mass margins. . . . . . . . . . . . . . 175
Figure 5.3 : 1000 kg PL vehicle stage mass margins after the use of composites. 176
Figure 5.4 : 1000 kg PL vehicle stage 2 mass margins and L/D ratios of

corresponding vehicle. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 178
Figure 5.5 : 500 kg payload vehicle L/D vs PMF. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 180

xxiv



Figure 5.6 : 500 kg payload vehicle stage 2 mass margin vs PMF. . . . . . . . . . . . . . . . . 181
Figure 5.7 : 500 kg payload vehicle stage 1 mass margin vs PMF. . . . . . . . . . . . . . . . . 181
Figure 5.8 : 500 kg payload vehicle mass vs PMF. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 182
Figure 5.9 : 500 kg payload vehicle paylaod ratio vs PMF. . . . . . . . . . . . . . . . . . . . . . . . . 183
Figure 5.10 :Variation of mass margin with respect to first stage T/W ratio and

number of engines. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 192
Figure 5.11 :Total thrust frame mass distribution as a function of T/W and

engine count. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 193
Figure 5.12 :Aggregate engine mass as a function of T/W ratio and number of

engines. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 193
Figure 5.13 :TVC system total mass dependence on engine count and T/W ratio. 194
Figure 5.14 :Burn time of the first stage for varying thrust-to-weight ratios. . . . . . . 195
Figure 5.15 :Azimuth range of Naro Space Center. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 200
Figure 5.16 :Azimuth range of Cape Canaveral. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201
Figure 5.17 :Azimuth range of Tanegashima Launch Center. . . . . . . . . . . . . . . . . . . . . . . 201
Figure 5.18 :Azimuth range of Northern Istanbul Launch Complex. . . . . . . . . . . . . . . 201
Figure 5.19 :Azimuth’s and corresponding inclinations of Naro Space Center. . . . 202
Figure 5.20 :Azimuth’s and corresponding inclinations of Cape Canaveral. . . . . . . 202
Figure 5.21 :Azimuth’s and corresponding inclinations of Tanegashima Launch

Center. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203
Figure 5.22 :Azimuth’s and corresponding inclinations of Northern Istanbul

Launch Complex. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203
Figure 5.23 :∆V profile of Naro Space Center. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204
Figure 5.24 :∆V profile of Cape Canaveral. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205
Figure 5.25 :∆V profile of Tanegashima / Takezaki Launch Center. . . . . . . . . . . . . . . . 205
Figure 5.26 :∆V profile of Northern Istanbul Launch Complex. . . . . . . . . . . . . . . . . . . . 206
Figure 5.27 :Launch mass of the configurations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209
Figure 5.28 :payload fractions of the configurations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210
Figure 5.29 :Configuration 1 engine numbers vs mass margin. . . . . . . . . . . . . . . . . . . . . 213
Figure 5.30 :Configuration 1 engine numbers vs structural mass. . . . . . . . . . . . . . . . . . . 213
Figure 5.31 :Configuration 2 engine numbers vs mass margin. . . . . . . . . . . . . . . . . . . . . 215
Figure 5.32 :Configuration 3 engine numbers vs mass margin. . . . . . . . . . . . . . . . . . . . . 216
Figure 5.33 :Configuration 4 stage 1 engine numbers vs mass margin. . . . . . . . . . . . . 217
Figure 5.34 :Configuration 4 stage 2 engine numbers vs mass margin. . . . . . . . . . . . . 217
Figure A.1 : Packing for n = 2 to n = 11. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 240
Figure B.1 : Mission case 1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 241
Figure B.2 : Mission case 2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 242
Figure B.3 : Mission case 3. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 242
Figure B.4 : Mission case 4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 243
Figure C.1 : Velocity benefit/penalty with launch latitude and azimuth.. . . . . . . . . . . 248

xxv



xxvi



PRELIMINARY DESIGN TOOL FOR LAUNCH VEHICLES
WITH LIQUID ROCKET ENGINES

SUMMARY

The design of launch vehicles powered by liquid rocket engines is a complex and
multidisciplinary challenge that requires careful integration of mission requirements,
propulsion system performance, and structural efficiency. Decisions made during
the preliminary design phase have significant impacts on the overall performance,
feasibility, and cost of the vehicle. Addressing the need for a structured and
efficient approach to this phase, this study introduces a computational tool designed
to streamline the early-stage design process of launch vehicles.

The tool operates through a systematic workflow, beginning with the calculation
of deltaV requirements for various mission scenarios. It evaluates the chemical
performance of different propellant combinations using a pre-generated data set,
providing rapid and accurate assessments essential for early design decisions. Users
are able to input key design parameters such as payload capacity, the number of stages,
and general engine performance characteristics.

For stage deltaV distribution, the tool offers flexibility by allowing either user-defined
input or an automated optimization process to minimize total vehicle mass. This
adaptability supports both custom mission-driven configurations and efficient design
solutions. The tool also performs a comprehensive scan of the vehicle’s diameter range
to ensure that mass margins are feasible. This process incorporates mass estimation
models for structural components and subsystems, supporting realistic and practical
design outcomes.

Additionally, the tool allows users to designate design concepts, including general
tank configurations and preferences, make sure that the resulting designs align with
structural and operational constraints. After determining the configuration, the
tool finalizes the vehicle’s dimensions, mass properties, and structural margins. It
further provides comprehensive cost analysis, estimating the costs of research and
development, production costs, and potential development timelines.

The tool’s accuracy and reliability have been validated using data from existing or
retired launch vehicles, demonstrating its effectiveness in producing feasible and
realistic design solutions. By integrating performance analysis, structural evaluation,
and cost estimation within a unified framework, the tool offers launch vehicle designers
a valuable resource for making informed decisions during the early stages of launch
vehicle development.

This study contributes to the advancement of preliminary design methodologies in
the space industry by providing a practical, flexible, and efficient approach to launch
vehicle configuration and analysis. The developed tool supports the creation of

xxvii



innovative and cost-effective launch systems, bridging the gap between preliminary
design and detailed engineering.

xxviii



SIVI ROKET MOTORLU FIRLATMA ARAÇLARI İÇİN
ÖN TASARIM ARACI

ÖZET

Sıvı yakıtlı çok kademeli fırlatma araçlarının tasarımı, görev gereksinimlerinin
mühendislik doğruluğuyla karşılanması, yapı-kütle dengelerinin sağlanması ve itki
sistemlerinin yeterli performansı sunması gibi çok sayıda alt problemi içeren karmaşık
bir mühendislik sürecidir. Özellikle ön tasarım aşamasında yapılan teknik kararlar,
daha sonraki detay mühendisliği ve maliyet analizlerini doğrudan etkilemekte;
sistemin uygulanabilirliği ve fizibilitesi açısından kritik roller üstlenmektedir. Bu
çalışma, bu ön tasarım aşamasında kullanılmak üzere, sistematik ve modüler yapıda
bir hesaplama aracı geliştirerek fırlatma aracı konfigürasyonlarının etkin şekilde
değerlendirilmesini ve analiz edilmesini amaçlamaktadır.

Geliştirilen hesaplama aracı, görev hedeflerinin belirlenmesiyle başlayıp kademeli
olarak araç geometrisinin, kütle dağılımının, yakıt konfigürasyonlarının ve performans
parametrelerinin optimize edilmesine olanak tanıyan bir süreç sunmaktadır. Araç, hem
kullanıcı tanımlı verilerle hem de otomatik optimizasyon modlarıyla çalışabilmekte,
bu yönüyle esnek bir tasarım ortamı sunmaktadır. Özellikle toplam araç kütlesini
minimize edecek şekilde kademeler arası Delta-V dağılımını otomatik olarak optimize
edebilmektedir. Bu işlevsellik, kullanıcıya hem görev odaklı özel çözümler üretme
hem de tasarım uzayında geniş bir aralıkta sistematik taramalar yapma imkanı
sağlamaktadır.

Araç, öncelikle görev senaryosuna uygun Delta-V ihtiyacını belirlemekle işe
başlar. Bu analiz, görev yörüngesinin irtifası, eğikliği ve fırlatma sahasının coğrafi
konumu gibi kriterlere bağlı olarak otomatik hesaplanır. Ardından, farklı yakıt
kombinasyonlarına ait termokimyasal verileri içeren önceden oluşturulmuş veri setleri
aracılığıyla motor performans parametreleri (özgül itki, çıkış hızı, genleşme oranı vb.)
elde edilir. Bu veriler, görev gereksinimlerine uygun motor tasarımı için başlangıç
noktası oluşturur.

İtki sistemi analizinde farklı çevrim türleri (örneğin gaz jeneratörlü ve aşamalı yanmalı
çevrimler) ve yakıt-oksitleyici çiftleri (özellikle LOX-RP1 ve LOX-LH2) dikkate
alınmakta, farklı genleşme oranları ve odak basınçları ile bu kombinasyonların motor
performansına etkisi analiz edilmektedir. Özellikle üst kademede kullanılan LOX-LH2
sistemleri, yüksek özgül itki avantajı sayesinde kütle optimizasyonuna önemli katkılar
sağlamaktadır. Her kademe için kullanıcı, motor karakteristiği, motor demeti sayısı ve
itki-ağırlık oranı gibi parametreleri girebilmekte veya belirli sınırlar dahilinde sistemin
bunları otomatik optimize etmesine olanak tanıyabilmektedir.

xxix



Araç içerisinde yer alan kütle tahmin modülü, her alt sistem (motorlar, tanklar,
yönlendirme sistemleri, ara bağlantı elemanları vb.) için birden fazla regresyon veya
ampirik model içermektedir. Bu modeller, geometrik parametreler, itki kapasitesi,
motor konfigürasyonu, yakıt hacmi gibi girdilere dayalı olarak yapısal ve kuru
kütle tahminleri üretmektedir. Tank kütlesi, genellikle çap, boy oranı ve içerdiği
yakıt hacmiyle ilişkili olarak tahmin edilirken; motor montaj sistemleri, yönlendirme
sistemleri ve ara bağlantı yapıları gibi bileşenlerde alan, hacim ve yük taşıma kapasitesi
gibi parametreler kullanılmaktadır. Bu modellerin hesaplama prensipleri genellikle
logaritmik, güç yasalarına dayalı veya çoklu değişkenli regresyonlar şeklindedir.
Aynı bileşen için farklı mühendislik yaklaşımlarının çıktıları karşılaştırılarak, tasarım
marjlarını sağlayan en güvenli tahmin tercih edilmektedir.

Bu kütle tahmin modeli, yalnızca geometrik ve fiziksel büyüklüklere dayanmamakta,
aynı zamanda yapısal marjlar ve kütle oranları gibi mühendislik toleranslarını da
dikkate almaktadır. Bu sayede elde edilen sonuçlar, yalnızca minimum kütleye
odaklanan bir optimizasyondan ziyade, uygulanabilirliği yüksek gerçekçi tasarımların
elde edilmesini sağlamaktadır.

Sistemde yer alan maliyet tahmin alt modülü, geliştirme, üretim ve operasyonel
maliyetleri ayrı ayrı analiz etmektedir. Bu analizde hem yıllık insan gücü gereksinimi
hem de üretim hacmine göre ölçeklenen değişken maliyetler dikkate alınmakta,
böylece farklı konfigürasyonlar arasındaki ekonomik karşılaştırma yapılabilmektedir.
Maliyet modeli ayrıca geliştirme süresine bağlı olarak toplam yatırım planını da
sunmaktadır.

Araç, farklı görev ve tasarım senaryolarını içeren üç vaka çalışmasıyla doğrulanmıştır.
Her biri, aracın alt modüllerinin koordineli biçimde çalışmasını sağlayarak farklı
teknik tercihlere göre nasıl sonuçlar ürettiğini göstermektedir.

İlk vaka çalışması, 500 kg ve 1000 kg faydalı yükleri LEO’ya taşıyabilecek iki farklı
küçük fırlatma aracı konfigürasyonu üzerine kurgulanmıştır. Görev için belirlenen
sabit Delta-V değeri kullanılarak, farklı yakıt oranları (örneğin propellant mass fraction
– PMF değerleri %0.86 ila %0.92 arasında) taranmış, çap-boy oranları ve motor demeti
sayıları optimize edilmiştir. 1000 kg taşıyan sistemde 9 motorlu bir ilk kademe ve
1 motorlu bir üst kademe tercih edilmiştir. Bu yapı, %1.46 faydalı yük oranı ve
%13’ün üzerinde yapısal marj üretmiştir. Daha küçük 500 kg konfigürasyonunda
ise PMF değeri düşürülerek %1’e yakın bir faydalı yük oranı korunmuş, yapısal
bütünlük sağlanmıştır. Bu analiz, aracın farklı yük profillerine adaptasyon kabiliyetini
ve optimum yapı-kararlarının nasıl değiştiğini göstermektedir.

İkinci vaka çalışması, ilk kademede kullanılan motor demeti sayısı (1’den 15’e) ve
itki-ağırlık oranının (1.1–1.8) aracın toplam kütlesine, itki yönlendirme sistemine,
motor kütlesine ve yanma süresine etkilerini incelemiştir. Bu analizle, optimum
performans bölgesi tanımlanmıştır: 6 ila 9 motorlu yapı, itki-ağırlık oranı 1.3–1.5
olan sistemler, hem yapısal hem operasyonel olarak dengeli sonuçlar vermiştir. Motor
arızasına dayanıklı, makul motor kütlesi ve kontrollü yanma süresine sahip sistemlerin
bu bölgede kümelendiği görülmüştür. Bu vaka, aracın parametrik tarama yeteneğini ve
çok boyutlu tasarım uzayında karar destek üretme kabiliyetini göstermektedir.

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Üçüncü vaka çalışması, 100 metrik tonluk bir faydalı yükün 800 km irtifada, 99 derece
eğiklikteki Güneş eşzamanlı bir yörüngeye ulaştırılmasını amaçlayan süper-ağır bir
fırlatma aracı tasarımıdır. Türkiye için önerilen İstanbul fırlatma kompleksiyle birlikte,
Tanegashima (Japonya), Naro (Güney Kore) ve Kennedy Space Center (ABD) sahaları
analiz edilmiştir. Japonya ve Kore’den yapılan fırlatmaların eğiklik farkının az olması
nedeniyle toplam Delta-V ihtiyacı yaklaşık 10150 m/s civarında kalmıştır. ABD’den
fırlatmada ise uçuş yönü sınırlamaları nedeniyle 14600 m/s’yi aşan çok daha yüksek
bir enerji ihtiyacı doğmuştur. Bu durum, fırlatma sahasının sistem kütlesi ve tasarım
üzerindeki etkisini açıkça ortaya koymuştur.

Bu vaka kapsamında altı farklı motor tipi; LOX-RP1 ve LOX-LH2 yakıt çiftleri,
farklı genleşme oranları ve çevrim tipleriyle birlikte değerlendirilmiştir. İlk kademede
yüksek itki sağlayan RP1 motorları, üst kademede ise yüksek özgül itki sağlayan LH2
motorları tercih edilmiştir. Yapısal marjları pozitif tutmak ve toplam sistem kütlesini
sınırlamak adına PMF değerleri %0.90 ila %0.95 arasında taranmış, motor performans
tabloları ile çap ve oran analizleri entegre edilmiştir.

Bu çalışma ile geliştirilen modüler hesaplama aracı, geleneksel basitleştirilmiş
analizlerin ötesine geçerek, her alt sistem için mühendislik doğruluğuna sahip
hesaplamalar gerçekleştiren, esnek ve uygulamaya dönük bir araç sunmaktadır.
Geliştirilen sistem, hem akademik çalışmalar hem de endüstriyel mühendislik için ön
tasarım ve detay mühendislik süreçleri arasında köprü kuran güçlü bir platform olma
niteliğindedir.

xxxi



xxxii



1. INTRODUCTION

In aerospace engineering, creating the ideal launch vehicle design has been a

top priority, and Multidisciplinary Design Optimization (MDO) techniques have

made a substantial contribution to this effort. To find the optimal vehicle design,

Castellini, Woodward, and others [1]–[5] have investigated combining structural

optimization, propulsion efficiency, and trajectory analysis into a single framework.

The complexity of the initial design phase, where precise mass estimation, Delta-V

(∆V ) budgeting, and subsystem-level analysis are crucial in defining workable and

effective configurations prior to full-scale optimization, is frequently overlooked by

MDO methodologies, despite their advancements.

The need for trustworthy engineering models that can strike a compromise between

computational efficiency and physical accuracy is a basic challenge in the early

design of launch vehicles. Although current MDO frameworks offer a comprehensive

approach to system-level optimization, they usually depend on streamlined mass

estimating methods and presumptions about vehicle subsystems. This may result

in early design errors that compromise the validity of optimization outcomes.

Furthermore, the ultimate cost, performance, and viability of a launch system are

greatly impacted by early design choices made in real-world engineering practice.

For this reason, it is essential to create computational tools that prioritize thorough

engineering modeling prior to optimization.

1.1 Purpose of Thesis

In response to these limitations, this research proposes a highly detailed and modular

computational framework tailored for the early-stage design and analysis of launch

vehicles. Rather than relying solely on an integrated MDO model to determine

an optimal configuration, this study segments the design process into multiple

specialized tools, each dedicated to a specific subsystem or analysis domain. By

1



adopting this approach, it becomes possible to capture key engineering constraints,

physical interactions, and performance trade-offs with greater accuracy, ensuring that

subsequent optimization studies are built upon a solid foundation of realistic and

well-defined parameters.

This thesis’s main goal is to create a comprehensive and adaptable design environment

that makes it easier to optimize launch vehicle designs iteratively. The framework

combines performance evaluation, structural mass estimation, and propulsion system

analysis into a unified but modular process. Before committing to a final optimization

procedure, this allows engineers and researchers to investigate potential design

trade-offs, carry out sensitivity assessments, and assess the effects of various

propulsion and structural configurations.

1.2 Similar Works in Literature

A strong methodological approach combining trajectory analysis, propulsion

system design, structural considerations, and cost evaluations is necessary for the

detailed design and optimization of launch vehicles. By creating computational

tools, suggesting optimization techniques, and examining current launch vehicle

configurations, numerous studies have made substantial contributions to this field. An

overview of related works in the literature is provided in this section, with an emphasis

on their approaches, conclusions, and limitations.

Numerous scholars have put forth frameworks, methods, and techniques over the

years to maximize the performance of both reusable and disposable launch vehicles.

Important contributions to the literature are reviewed in this section. The review seeks

to shed light on conceptual design tools, cost-reliability trade-offs, and optimization

techniques that support launch vehicle development in its early stages.

The modeling and design processes of launch vehicles have been extensively studied

in the literature. Some of these studies include Multidisciplinary Design Optimization

(MDO) techniques, while others focus solely on specific subsystem modeling. This

section explores the key works in MDO applications for launch vehicle design, mass

estimation methods, and hybrid propulsion systems.

2



The first subject to be discussed is the work of Castellini, which is a creation of a

framework for Multidisciplinary Design Optimization (MDO) of expandable launch

vehicles [1]. The mentioned work is one of the detailed studies done on a launch

vehicle design, which incorporates not only the bigger picture, but also detailed

sub-system level modelling. The primary motivation behind this research is the high

time and cost investment in the preliminary and conceptual design phases. Castellini’s

study claims that 80% of the total Life Cycle Cost (LCC) of a launch system is

determined during these early design stages. His goal is to retain a manageable

computational effort while increasing the multidisciplinary models’ accuracy and

robustness in comparison to conventional fixed-point iteration design techniques. For

the development and validation of the MDO framework he went into a 2-step approach

where at first he created a conceptual level model, he calls V1. Following that, these

models were verified using current European rockets like VEGA and Ariane 5. The

validation revealed a number of flaws, which prompted the creation of more thorough

early preliminary (V2) models. The accuracy of the V2 models was significantly

higher than that of the V1 models, with errors on payload mass estimation for Ariane

5 and VEGA decreasing from about 10% to about 4-6%. With nearly 5% system-level

mistakes resulting from average errors on an individual mass figure, worst-case

sensitivity tests showed that the total inert masses continue to be the most crucial

figures, where the error of 3% for propulsion performances and 1% for aerodynamic

coefficients observed.

Woodward addressed a gap in conceptual launch vehicle design software. As claimed

by Woodward, his work fills this gap by creating a modular launch vehicle design

tool that is based on Python and supports iterative revisions, making it a flexible tool

for launch vehicle analysis and optimization [2]. In order to facilitate the conceptual

and preliminary design of launch vehicles, the developed launch vehicle design tool

was outfitted with a number of essential features. The energy requirement calculation,

which establishes the total energy required for the vehicle to reach a specific orbit with

a specified payload capacity, was one of the tool’s main features. This is essential

for specifying the needs for staging and propulsion. Mass estimate, which determines

the masses of various parts to evaluate the structural and propellant mass distribution

3



and guarantee a well-balanced and optimized vehicle architecture, was another crucial

aspect. In addition to this, propulsion system sizing was used to ensure that the system

satisfies the thrust and specific impulse requirements of the mission, which is crucial

for reaching the necessary Delta-V (∆V ). The program also included calculations for

propellant mass and volume, which allowed for precise assessment of the required fuel

and oxidizer volumes while optimizing tank layouts and dimensions. Additionally,

geometric modeling was incorporated, enabling the launch vehicle’s basic dimensions

and form while preserving a useful and effective design. The entire ascent-to-orbit

flight route could be modeled using a trajectory modeling module, which included

aerodynamic effects to precisely forecast performance in the actual world. In order

to guarantee that the vehicle design stays aerodynamically stable and controllable

throughout the ascent phase, a basic stability and control study was also included. Last

but not least, the program included a cost estimation function that made it possible to

project development and manufacturing expenses as well as the launch service price

required for commercial viability.

There is another study by Ballard for the development of similar work. Ballard’s study

is concentrated on the preliminary design and fixed-point optimization techniques for

the optimization of launch vehicles [3]. She created a computational tool in MATLAB.

The staging optimization and gravity turn trajectory refinement were made possible

by this tool’s integration of an optimization framework based on the Nelder-Mead

algorithm. Improving an already-existing stage design tool that was created in Excel

was a crucial component of the study. In order to assess crucial propellant budgets,

helium reserves, and structural mass estimates, this instrument was greatly enhanced.

A more precise comprehension of stage-level performance indicators was ensured by

the improved tool’s more accurate launcher feasibility assessments.

Another study that deals with the multidisciplinary design optimization (MDO) of

launch vehicles is that of Balesdent [4]. The extremely limited and computationally

demanding nature of trajectory optimization, which has a major influence on the entire

design process, is a major problem in this field. The work presents a revolutionary

Stage-Wise Decomposition for Optimal Rocket Design (SWORD) methodology in

order to overcome this. In contrast to conventional MDO techniques, SWORD

4



breaks down the optimization problem into several flight phases, handling each one

separately while coordinating its optimizations. By breaking down the otherwise

complex multistage launch vehicle design problem into a number of easier-to-manage

subproblems, this organized approach increases computational efficiency. The study

examines the optimization of a three-stage-to-orbit launch vehicle by contrasting

the SWORD method with the traditional Multi-Discipline Feasible (MDF) MDO

approach. According to the results, SWORD improves optimization efficiency,

especially when it comes to finding workable solutions more quickly and producing

better optimal designs in a limited amount of computational time. To further improve

convergence time, an optimization technique specifically designed for SWORD was

created. This enables the method to function efficiently without the need for manual

initialization parameters or prior knowledge of the search space. This development

in launch vehicle design optimization is a significant addition to the field of space

transportation system design since it offers a viable strategy for reaching quicker and

more dependable solutions.

Gaspar’s research is another contribution to launch vehicle design studies, developing

a MATLAB-based tool that integrates mass modeling and trajectory simulation for

early-stage rocket design [5]. Using a database to define reasonable parameter

ranges, the tool estimated key vehicle parameters such as length, diameter, thrust,

burning duration, and specific masses. While Mass Estimating Relationships

(MERs) were attempted, reliable structural estimations were limited due to data

scarcity. The mass model iteratively applied Tsiolkovsky’s equation and heuristic

MERs to ensure realistic structural mass predictions, incorporating geometric

computations to refine accuracy. The trajectory simulation included three phases:

vertical ascent, gravity turn (determined by the Knudsen number), and free-flight,

optimizing time and propellant usage. A simplified drag model with Mach-dependent

coefficients and a hybrid atmospheric model (combining 1962 and 1976 US standard

atmospheres) were employed. The tool was validated using the Proton K and

Vega launch vehicles, demonstrating effective trajectory simulations but encountering

discrepancies due to idealized assumptions such as constant thrust and full propellant

consumption. Structural mass estimates were generally high, while propellant

5



mass was underestimated, resulting in lower-than-expected Gross Lift-Off Weight

(GLOW). Additionally, structural margin considerations were lacking, limiting its

preliminary design accuracy. After validation, the tool was applied to an Ariane

5-based rocket design, successfully reducing GLOW while maintaining mission

objectives. Optimization factors included thrust levels, Delta-V (∆V ) distribution,

and diameter selection, with boosters enhancing payload capacity. Despite its

limitations in structural estimation, the tool provides a useful framework for early-stage

launch vehicle design, offering insights into performance trade-offs and enabling

configuration optimization.

1.3 Hypothesis

Developing an ideal launch vehicle configuration using Multidisciplinary Design

Optimization (MDO) is the main focus of the literature now available on launch vehicle

design, especially the works of Castellini, Woodward, Gaspar, Balesdent, and others.

By combining trajectory analysis, propulsion efficiency, and structural optimization

into a single framework, these studies seek to determine the optimal vehicle design.

Nevertheless, this method frequently ignores the complexities of the initial design

stage, where precise mass estimation, Delta-V (∆V ) budgeting, and subsystem-level

analysis are essential prior to optimization.

This thesis, on the other hand, aims to create a highly comprehensive and flexible

computational framework that acts as a fundamental tool for early-stage design

and optimization studies, rather than simply determining the ideal launch vehicle

configuration. This research divides the launch vehicle design process into several

specialized tools, each of which is in charge of a particular subsystem or analysis

domain, rather than using a single integrated MDO model. This makes it possible to

comprehend the engineering and physical limitations that need to be taken into account

before a thorough optimization study can be carried out in greater detail and with

greater accuracy.

6



1.4 Thesis Outline

Chapter 1 explains the study’s motivation and the difficulties in designing launch

vehicles, especially in the beginning. A discussion of the shortcomings of current

Multidisciplinary Design Optimization (MDO) techniques emphasizes the need for

a more precise and adaptable computational framework. A synopsis of the thesis

contributions is provided, along with the goals and parameters of the study.

Chapter 2 offers a thorough analysis of current launch vehicle design studies and

methodology, giving readers a basic understanding of the area. The historical

development of launch vehicles is examined first, following developments from the

earliest rocketry to the creation of contemporary reusable systems. This historical

perspective draws attention to the technological turning points that have influenced

modern design methodologies. The conversation then shifts to new developments

in launch vehicle design, emphasizing inventions like reusable launch systems and

specialized tiny satellite launchers that are propelling space access in the future.

Chapter 3 describes how the computational tool was developed, including its modular

structure and how its sub-models interact with one another. It discusses the

propulsion system model, which computes propellant mass, oxidizer-to-fuel ratio,

thrust, and specific impulse. The mass estimation model, which predicts component

mass distribution using empirical data and Mass Estimating Relationships (MERs),

is discussed. Additionally included is structural modeling, which takes into account

subsystem contributions, tank dimensions, and fairing mass. The cost estimating

approach, which evaluates development, production, and operating expenses to

determine the economic viability of various configurations, is also explained in this

chapter.

Chapter 4 contains the validation. In this chapter, the correctness and dependability of

the computational tool’s sub-models are evaluated in order to validate it. To make sure

each model is consistent with the characteristics of launch vehicles in the actual world,

it is validated against reference scenarios and current data. Benchmark comparisons

and existing empirical data are used to assess the propulsion system, mass estimation,

7



structural, and cost estimation models. In order to pinpoint possible error sources and

improve the tool’s predictive power, discrepancies are examined.

Chapter 5 includes case studies, in order to illustrate the usefulness of the

computational tool in launch vehicle design. To evaluate the tool’s performance under

various mission profiles and design limitations, a variety of vehicle configurations

are examined. The case studies investigate the effects of design decisions on overall

vehicle performance and viability, including mass distribution, structural factors, and

propulsion selection. This chapter also examines the outcomes of the case studies and

evaluates how well the computational tool performs in producing dependable launch

vehicle designs. The accuracy and consistency of the sub-models are evaluated by

examining key findings within the framework of current techniques. The effect of

design decisions on the overall performance of the vehicle is investigated, paying

special emphasis to cost prediction, propulsion efficiency, and mass estimation

accuracy.

Chapter 6 summarizes the study’s general findings and highlights the usefulness

and contributions of the created computational tool for launch vehicle design. The

research’s weaknesses are noted and highlighted, indicating areas that require more

development. In order to further improve the design process, prospective future

development directions are finally described, including improvements to current

models and potential integration with more sophisticated optimization frameworks.

8



2. FUNDAMENTALS AND EVOLUTION OF LAUNCH VEHICLE DESIGN

2.1 Launch Vehicles: Historical Perspective and Current Trends

For thousands of years, humans have been amazed by the stars and have longed to

travel across the sky and discover the unknown. A result of creativity, science, and

willpower, space travel has revolutionized exploration and set precedents that have

permanently altered our perception of the cosmos and our place in it.

The advancement of space flight is a tribute to human intellect and tenacity as well as a

technological success. In order to build machines that could resist the harsh conditions

of space, it required advances in physics, materials science, computer engineering,

and propulsion systems, all of which had to be seamlessly integrated. Space flight is

a prime example of what humanity can accomplish when desire and ingenuity are in

harmony, from the initial hesitant steps of putting satellites into orbit to the landing

of humans on the moon and the sending of probes to the extreme reaches of the solar

system.

As a peak of technical development, space travel has not only deepened our

understanding of the universe but also sparked numerous innovations that have

improved life on Earth, ranging from global positioning systems (GPS) and satellite

communications to breakthroughs in materials and medicine. Space travel is more

than just a technological achievement; it is a representation of our insatiable curiosity

and our common goal of expanding the realm of the possible. It serves as a constant

reminder that people can literally and figuratively aim for the stars.

History of early rocketry starts with Konstantin Tsiolkovsky. Many people consider the

Russian schoolteacher and self-taught scientist Konstantin Tsiolkovsky (1857–1935)

to be the founder of theoretical astronautics. Tsiolkovsky made important contributions

to our understanding of how rockets could be used to journey to space, despite working

in relative solitude. In 1903, he published his contributions to rocketry in the magazine

9



called Nauchnoe Obozrenie (which in Russian, it means "Scientific Review") with a

name of "Exploration of the World Space with Reaction Machines", which can be

observed from Figure 2.1 [6]. In his work that is mentioned, he gave the details of

his famous equation that defines the relationship between the velocity of a rocket, the

exhaust velocity of its propellant, and the mass ratio of the rocket before and after fuel

consumption, Equation 2.1.

∆v = ve ln
(

m0

m f

)
(2.1)

Figure 2.1 : Tsiolkovsky’s work, 1914 edition.

While the Tsiolkovsky’s work is entirely theoretical, in order to reach the stars, one

must use theory to create the practical knowledge. That is the text-book definition

of Robert Goddard. Goddard’s work has been more of an experimental kind and he

is called "The Father of Modern Rocketry", because of his contributions to modern

rocketry as he had a large number of inventions and patents, exceeding 200 [7]. One

of the contributions of his work to rocketry is the world’s first liquid propellant rocket

engine and rocket, which can be seen from Figure 2.2 [6].

10



Figure 2.2 : Left. Goddard’s liquid propellant rocket. Right. Goddard’s rocket’s
components.

The rocket that is observed in Figure 2.2, is launched in 16 March 1926. As far as

it can be understood, he couldn’t find a proper place to launch. Thus, he set the

launch place as his aunt’s (Ms. Effie M. Ward) backyard at the farm located in Auburn,

Massachusets. He managed to launch the rocket to the altitude of 12.5 m and landed

56 m far from the launch location in a cabbage field [7]. The flight took a little over 2

seconds, but became a milestone in modern rocketry. He was also known for his works

for the war efforts of World War II. His contributions to the rocketry are recognized and

has been given a task of rocket-assisted takeoff development for heavily loaded aircraft

operations. Which in fact, tested and approved. But unfortunately, he was not enough

recognized by the scientific community until his passing in 1945. He was awarded

with Congressional Gold Medal, 14 years after his death and recognized as "father of

modern rocketry". In the same year his name is given to a space center which was

formerly known as Beltsville Space Center, renamed into Goddard Space Center [6].

In the times of World War II, Germany made efforts to use the technology of

rocketry in its war efforts. The team led by Dr. Wernher von Braun developed the

Nazi-Germany’s well-known guided ballistic rocket [6] called V-2 (Vergeltungswaffe

2, Eng. Retribution Weapon 2) which also called with the name of Aggregat 4 (A-4),

can be examined from Figure 2.3 [7].

11



Figure 2.3 : V-2 ballistic missile.

When the V-2 was tested, its first 3 launch attempts were unsuccessful. But the fourth

time it accomplished a maximum altitude of 96 km and a range of 193 km from the

launch site. Other important specifications of V-2 is given in Table 2.1 [7]. With that

achievement of V-2, the leader of the development team, Dr. von Braun, was appointed

professor at that time [7].

Table 2.1 : V-2 missile specifications.

Parameter Value
Takeoff mass [kg] 12700
Warhead mass [kg] 1000

Total length [m] 14
Liftoff T/W 2
ECO T/W 7

12



The liquid propellant rocket engine that is developed for V-2 missile was the most

powerful for its time. Specifications of the engine of V-2 can be observed from Table

2.2 [7]. Also, the photograph of engine can be observed from Figure 2.4 [7].

Table 2.2 : V-2 missile liquid engine specifications.

Parameter Value
Thrust @SL [kN] 245

Thrust @ECO [kN] 311.3
Chamber pressure [bar] 15.2

Chamber temperature [◦C] 2300
Throat diameter [mm] 400
Exit diameter [mm] 740

Exhaust velocity [m/s] 2050
Specific Impulse (Isp) @SL [s] 210

Figure 2.4 : V-2 engine.

In the last minutes of World War II, it was obvious for both Soviet Union and Allies

that, rocketry would be the milestone of technological advancement. Therefore, this

was the reason for the both sides to collect assets from Nazi-Germany, some of them

are individuals. There was an operation called "Operation Paperclip", military of

United States of America, captured scientists including and led by Dr. Wernher von

Braun. Soviet Union captured some other scientists as well. Those progresses were

the cornerstone of the rocket technology [6].

13



With the knowledge they gathered and experience they made, both USA and Soviet

Union developed ballistic missiles for foreign threat. Both also realized that those

ballistic missiles could be used as a space launch vehicle to launch objects into orbit.

Soviet Union was the first country to launch a manmade object into orbit. The

calender’s were October 1957 and the launched object was Sputnik. The Sputnik was a

simple satellite with a simple mission but it was a crucial milestone which also started

the space-race between United States and Soviet Union.

Figure 2.5 : R-7 Semyorka to Soyuz launch vehicle.

The Sputnik was launched into orbit with R-7 Semyorka ICBM based launch vehicle

[7]. R-7 vehicle used extensively for launching the objects into space for Soviet space

program and it turned into the Soyuz, which is still in use today by Russia. This

evolution can be observed from Figure 2.5 [6]. Although Soviet’s have the R-7 ICBM,

its launch capacity was not enough for launching moon mission payloads. Therefore,

they aimed to create a gigantic launcher called N-1, which was also the counterpart of

Saturn V. The magnitude of N-1 can be seen from the Figure 2.6. N-1 had a total of 42

rocket engines, 30 in its first stage, 9 in its second stage and 4 in its third stage [8].

While Soviet’s were struggling with harsh development and flight test phases of N-1

launch vehicle, United States was testing different concept vehicles to create a final

vehicle to land on the moon. Therefore, their attempts were successful. Saturn V with

code-name SA-501 was the vehicle that landed United States on the moon in 1969.

Saturn rocket family evolution can be observed from Figure 2.7 [6].

14



Figure 2.6 : Side and bottom views of N-1.

Saturn V had 5 F-1 engines on its first stage, 5 J-2 engines on its second stage, and 1

J-2 engine on its third stage. Other details of the Saturn V can be seen from Figure

2.8 [9].

15



Figure 2.7 : Family of Saturn rockets.

Figure 2.8 : Saturn V.

Saturn V was a part of the program Apollo, which started in 1962 and ended in 1972.

Saturn V later used for a mission, Skylab space station launch, and it became its last

flight before retirement. USA won the lunar part of the space race.

16



In the following years, the sector has evolved into space stations, Earth-orbiting

satellites, and orbiter technologies. For this effort, the US created the Space Shuttle [8].

Space Shuttle launch can be observed from Figure 2.9 [6]. The Soviets also had an

orbiter, so called Buran, which looks alike the Space Shuttle. Buran orbiter has been

presented in Figure 2.10 [6].

Figure 2.9 : Space Shuttle launch.

Main difference of Shuttle and Buran is the configuration of the vehicle. Buran orbiter

is just a payload of Energia launch vehicle, and the launch vehicle can be used to

launch other payloads than the orbiter, while the Shuttle is just part of the launch

vehicle. Whereas the Shuttle is used broadly in US space missions until 2011, Buran

has just seen one orbital flight before its cancellation in 1994 [6].

Figure 2.10 : Buran orbiter.

17



Despite the technology that Shuttle had and offered, it demonstrated the operational

and financial difficulties of launch vehicle development and operations. As Launius

stated in his work, the cost for Shuttle Program was 70 K USD per kg (adjusted for

inflation to 2024 USD) [10]. Aware of the problem, US Senate launched an act in

1984 that enabled private companies to develop, launch and operate launch vehicles

[11]. This was a landmark for space launch sector, as it diminishes the monopoly of

governmental programs to create a foundation for the privatization of space operations.

This set of regulations promoted competition and gave private businesses the freedom

to design launch systems on their own. The act’s long-term effects were apparent in the

1990s and 2000s when businesses like SpaceX, Blue Origin, and Rocket Lab started

using technological breakthroughs to lower prices and boost the number of launches.

Figure 2.11 : Launch costs.

The development of more effective commercial space launch vehicles and rocket

technological developments are highlighted in Figure 2.11 [12], which shows the sharp

decline in launch costs over time. According to the statistics, the cost of launching

payloads to low Earth orbit (LEO) has dropped dramatically for heavy, medium,

and small launch vehicle categories between the 1960s and the 2020s. Costs were

historically very high, especially in the early years of space exploration when the sector

was controlled by government-sponsored missions. A major turning point was the

development of reusable launch vehicles, like SpaceX’s Falcon 9 and Falcon Heavy,

which allowed costs per kilogram to fall to previously unheard-of levels [12].

18



On the other hand, as the Figure 2.11 illustrates, small launch vehicles have a clear

lower-cost limit, but they have a clear benefit in the shape of missions that are

specifically designed for them. For customers that need unique orbital insertions or

exclusive usage of the rocket, small launchers offer customized options in contrast to

medium, heavy, or super-heavy launch vehicles, which usually carry many payloads

for cost-sharing. They are therefore the best option for satellite operators who value

accuracy and adaptability over financial savings. This advantage is pricey, too, because

small launch vehicles are more expensive per kilogram and are unlikely to be able to

meet the lower costs that larger vehicles can achieve.

Nonetheless, because of economies of scale, utilizing medium or heavy-lift vehicles

for smaller payloads can drastically lower launch costs; nonetheless, these shared

missions sometimes lack the accuracy and adaptability that specialized launches

provide. Operators may have to give up their desired orbit or timing if they decide

to take this approach. This trade-off demonstrates how small and large launch vehicles

play complementary roles in the market: small rockets serve high-priority, specialty

missions, while larger rockets maximize cost-effectiveness for aggregated payloads or

bulk deployments. When choosing a launch solution based on mission-specific needs

and financial constraints, customers must carefully consider these variables.

Looking toward the future, the economics of the launch industry is expected to

change further. Due to technological advancements, re-usability, and the growing

influence of commercial players, the launch industry is expected to experience

previously unheard-of cost reductions in the future. Current projections indicate

that advancements in re-usability, economies of scale, more effective production

techniques, and reduced input costs might drive a 95% reduction in launch costs by

2040, reaching $100 per kilogram [12].

19



2.2 Launch Vehicle Fundamentals

A launch vehicle is a system used for carrying cargo beyond Earth’s atmosphere. This

cargo includes satellites, scientific instruments, crewed spacecraft, and specific cargo

to other spacecraft (i.e. ISS). In order to overcome Earth’s gravity and deliver cargo

to their intended orbits or interplanetary paths, these spacecraft produce enormous

amounts of thrust. The development of commercial satellite deployment, space

exploration, and national security missions all depends heavily on the effectiveness,

dependability, and affordability of launch vehicles.

2.2.1 Launch vehicle classification

Depending on their design and the goals of their missions, launch vehicles have many

kinds of uses. Some are designed for satellite placements, putting Earth observation,

communication, or navigation satellites in geostationary orbit (GEO) or low Earth

orbit (LEO). Others send cargo outside of Earth’s orbit to support scientific missions

like space telescopes or interplanetary probes. Furthermore, cargo resupply missions

transport vital supplies to space stations such as the International Space Station

(ISS), while human spaceflight missions necessitate specific launch vehicles that can

guarantee crew safety. The goal of the ongoing development of launch vehicles is

to make space travel more economically and environmentally feasible by increasing

payload capacity, improving reusability, and lowering costs.

Traditionally, launch vehicles are categorized based on their payload capacity

(generally to LEO) and mission type. Classification made by NASA in one of their

technological road-map, is made with the capacity to low-earth-orbit (LEO), details of

this classification summarized in Table 2.3 [13].

The classification of launch vehicles based on mission types can be considered

somewhat arbitrary because the same launch vehicle can serve multiple mission

types depending on the payload and mission objectives. For example, a vehicle like

Falcon 9 can launch communication satellites [14], deploy cargo resupply missions

20



Table 2.3 : Launch vehicle classification by payload capacity.

Launch Vehicle Class Payload Capacity (mT)
Small Lift < 2

Medium Lift 2 - 20
Heavy Lift 20 - 50

Super-Heavy Lift > 50

to the ISS [15], and carry crewed missions using the Crew Dragon capsule [16].

Similarly, the Atlas V rocket has been used for Earth observation satellites [17],

interplanetary exploration (e.g., Perseverance rover [18]), and other mission types.

This overlap highlights that mission type classifications often depend more on the

specific payload or target orbit rather than the design of the launch vehicle itself.

Thus, the boundaries between mission types are fluid and not strictly defined. For

an overview of mission-based classifications, refer to Table 2.4.

Table 2.4 : Launch vehicle classification by mission types and their applications.

Mission Type Orbit/Target Examples
Satellite Deploy-
ments

LEO (300-2,000 km) Starlink, OneWeb (Communica-
tions)

MEO (2,000-35,786
km)

GPS, Galileo (Navigation)

GEO (35,786 km) Weather and Telecommunications
Satellites

SSO (500-800 km) Earth Observation, Remote Sensing
Crewed
Spaceflight
Missions

LEO ISS crew transport (Crew Dragon,
Soyuz)

Beyond Earth Artemis (SLS), Future Mars Mis-
sions

Cargo Resupply
Missions

LEO ISS resupply (Cargo Dragon,
Cygnus, Progress)

Interplanetary Lunar Artemis (SLS), Chang’e missions
Mars Perseverance (Atlas V), Tianwen-1

(Long March 5)
Asteroids/Outer Planets Europa Clipper, Lucy, JWST (Ari-

ane 5)

21



2.2.2 Components of launch vehicles

Modern launch vehicles are highly sophisticated engineering systems, meticulously

designed to transport payloads beyond Earth’s atmosphere. Every part is essential

to the mission’s success. All launch vehicles have the same main goal, which is to

deliver payloads into space, but their internal architecture differs based on propulsion

technologies, mission requirements, and reusability considerations. Understanding

the fundamental components of a launch vehicle is essential for understanding the

intricacies of spaceflight and the engineering challenges involved.

Many of the basic components of launch vehicles are almost the same, even

though they are made to fulfill different mission requirements. Certain fundamental

components, such as propulsion systems, avionics, and structural arrangements, are

present in all rocket designs, regardless of whether they are intended for the launch of

satellites, crewed missions, or deep-space probes. The Figure 2.12 gives a summary

of the main parts of a two-stage launch vehicle (that does not have solid rocket

motors) and a rundown of its crucial subsystems. This representation helps illustrate

how various elements work together to ensure a successful launch, from liftoff to

payload deployment. The basic architecture of launch vehicles adheres to a common

engineering framework, even though specific design decisions may differ depending

on factors like payload capacity, reusability, and propulsion type. Component details

of launch vehicle’s are presented in Table 2.5, 2.6, 2.7 and 2.8. It should be

noted that the components listed in the tables and illustrated in the figure do not

represent a comprehensive list of all possible launch vehicle subsystems. Instead,

they are selected to provide a general overview of key elements that contribute to the

vehicle’s overall function. Depending on the specific design, mission requirements,

and technological advancements, additional components may be incorporated, while

some listed elements may vary in configuration or implementation.

22



Figure 2.12 : Launch vehicle components.

23



Table 2.5 : Launch vehicle components (1 to 7) continued.

# Component
Name

Component Function

1 Payload Fairing Protects payload during ascent and is jettisoned after
reaching space

2 Payload Launch vehicle’s precious burden

3 PLA/PAF The adapter that connects the launch vehicle and
payload

4 2nd Stage Fwd.
Skirt

Structural element that connects the second stage and
payload related structures (PLA, Payload and Fairing)

5 2nd Stage Ox.
Tank

Oxidizer storage unit for second stage propulsion
system

6 2nd Stage Fuel
Tank

Fuel storage unit for second stage propulsion system

7 2nd Stage Eng.
Thrust Frame

Structural element that undertakes the load of second
stage rocket engine

Table 2.6 : Launch vehicle components (8 to 14) continued.

# Component
Name

Component Function

8 2nd Stage Eng.
TVC System

Adjusts the direction of the engine nozzle to precisely
control the vehicle’s orientation and trajectory during
orbital insertion and space maneuvers

9 1st Stage
Forward Skirt

Structural element that connects 1st stage upper tank
and interstage structure

10 1st Stage Ox.
Tank

Oxidizer storage unit for first stage propulsion system

11 Intertank
structure

Structural element that connects oxidizer and fuel
tank

12 1st Stage Fuel
Tank

Fuel storage unit for first stage propulsion system

13 1st Stage Eng.
Thrust Frame

Structural element that undertakes the load of first
stage rocket engine

14 Aerodynamic
Control Surfaces

Control element that is used for controlling the ve-
hicle’s orientation and trajectory during atmospheric
flight through aerodynamic forces

24



Table 2.7 : Launch vehicle components (15 to 21) continued.

# Component
Name

Component Function

15 1st Stage Engines Provides necessary initial thrust to lift the vehicle off
the ground and through the lower atmosphere.

16 1st Stage Eng.
TVC System

Dynamically tilts the main engines to steer the rocket
during ascent, providing stability, guidance, and
trajectory corrections to ensure a controlled flight path
through the atmosphere.

17 Cryogenic Gas
Pressurization
System

Propellant pressurization system that uses cryogeni-
cally stored inert gas. The pressurization gas is
stored in cryogenic conditions to increase the storage
density. Gas is fed into rocket engines heat exchanger,
where it heats up to a certain temperature. After that
it is fed to the propellant tank.

18 Propellant
feed-line through
tank

Feed-line of a propellant that goes through the other
propellant tank. Aerodynamic heating and residual
propellant mass is reduced relative to out-of-body
feed-line.

19 Ambient Gas
Pressurization
System

Propellant pressurization system that uses ambient
stored inert gas. Easy to manage relative to
cryogenically stored gas.

20 1st Stage Avion-
ics

System consists of flight computers, inertial mea-
surement units (IMUs), telemetry transmitters, and
communication systems, responsible for navigation,
thrust vector control (TVC), and stage separation
sequencing while ensuring continuous data transmis-
sion to ground stations throughout ascent.

21 2nd Stage
Engine(s)

Provides the necessary propulsion for orbital insertion
and payload deployment, operating in the vacuum of
space with a high-efficiency nozzle, thrust vector con-
trol (TVC) for precise maneuvering, and optimized
propellant consumption to achieve mission-specific
trajectories.

25



Table 2.8 : Launch vehicle components (22 to 28) continued.

# Component
Name

Component Function

22 Interstage Struc-
ture

Structural element that connects the first and second
stages.

23 Antislosh Baffles Internal structures within propellant tanks designed
to reduce fuel movement (sloshing) during flight,
ensuring stable mass distribution, preventing shifts in
the vehicle’s center of gravity, and improving engine
performance and flight stability.

24 Propellant Man-
agement Devices
(PMDs)

Internal components within propellant tanks that
ensure consistent fuel and oxidizer flow to the
engines by managing fluid movement, preventing
bubbles or vapor ingestion, and enabling reliable
engine performance in both gravity and microgravity
environments.

25 Autogenous Pres-
surization System

Pressurization system that uses heated gaseous
propellant from the engine cycle to self-pressurize
the propellant tanks, eliminating the need for external
pressurant gases like helium.

26 Insulation Components that minimizes heat transfer and pro-
tect critical components from extreme temperature
variations experienced during propellant storage,
ascent, and space exposure. It includes thermal
blankets, foam insulation, and multilayer insulation
(MLI) to prevent cryogenic propellant boil-off, reduce
aerodynamic heating, and shield avionics from
thermal fluctuations.

27 Tank Gas
Diffuser

Internal component within a propellant tank that
evenly distributes pressurizing gas, preventing lo-
calized turbulence, reducing bubble formation, and
ensuring uniform pressure stabilization to maintain
consistent propellant flow to the engines during flight.

28 2nd Stage Avion-
ics

System includes onboard processors, inertial mea-
surement units (IMUs), telemetry transmitters, and
communication systems, managing precise guidance,
thrust vector control (TVC), and payload deploy-
ment operations while maintaining continuous data
exchange for mission tracking and orbital corrections.

26



2.2.3 Launch vehicle mass elements

All components in a launch vehicle can be classified based on their function and

whether they remain part of the vehicle after propulsion or are utilized during the

flight. These elements are typically divided into three primary categories:

• Propellant Mass (mp): The propellant is the mass expelled through the vehicle’s

engines to generate thrust and propel the rocket forward. This category

encompasses all consumable elements that are ejected as exhaust gases through the

propulsion system. The total propellant mass includes the oxidizer and fuel, which

are consumed during flight to produce the necessary force for ascent. However,

non-propulsive jettisoned elements, such as payload fairings, are not included in

this category.

• Payload Mass (mPL): The primary objective of any launch vehicle is to transport a

payload to a designated orbit or trajectory. The payload mass includes any scientific

instruments, satellites, crewed modules, or cargo that the launch vehicle is designed

to deliver. Every other component of the rocket serves the purpose of ensuring

that the payload reaches the required velocity and altitude. Payload mass is often

protected by a fairing, which is typically discarded once the vehicle reaches the

upper atmosphere, thereby reducing unnecessary weight and improving efficiency

• Structural Mass (ms): The structural mass encompasses all non-propellant and

non-payload elements that remain with the vehicle during flight. This includes

the outer structure, tanks, rocket engines, avionics, plumbing/feed lines, insulation,

wiring, and guidance systems. While referred to as "structure," this mass also

accounts for unused residual propellant, which remains in the tanks after the burn

is completed. The structural mass, along with the payload, forms the final mass of

the vehicle once all usable propellant has been expended. Some elements, such as

payload fairings, are jettisoned mid-flight and do not contribute to the final mass.

However, all core components that do not detach during ascent remain part of the

structure until stage separation or final payload deployment.

27



2.2.4 Launch vehicle performance

One important determinant of a launch vehicle’s ability to transport payloads into

different orbits or outside of Earth’s gravitational pull is its performance. There are

a couple of factors that affect the performance of launch vehicles. Those factors are

defined as:

• Delta-V (∆V )

• Specific impulse (Isp)

• Payload fraction (π)

• Structural factor (σ )

Engineers can optimize launch vehicle design for cost-effectiveness, dependability,

and mission flexibility by being aware of these factors. Those main factors influencing

launch vehicle performance are examined in this section.

2.2.4.1 Delta-V requirements

As the name suggests, it is the velocity change needed for certain mission scenario.

The required velocity change (∆V ) for a rocket stage is given by the "Tsiolkovsky

Rocket Equation", which was also presented in Equation 2.1. In this instance it is

presented with the values of the rocket.

∆V = Isp ·g0 · ln
(

m0

m f

)
(2.2)

where:

• Isp = specific impulse (s),

• g0 = standard gravitational acceleration (9.8066 m/s2),

• m0 = initial mass (including propellant),

• m f = final mass (after propellant is burned).

28



Each space mission has a unique budget of energy (Delta-V budget), which solely

depends on the orbit or celestial body that is targeted. Figure 2.13 [19]visually

compares the Delta-V requirements for different missions. Table 2.9 gives an insight

about some example space mission energy requirements.

Table 2.9 : Typical Delta-V requirements for various space missions.

Mission Delta-V Requirement (km/s)
Low Earth Orbit (LEO) 9.3 - 9.8
Geostationary Transfer Orbit (GTO) 10.5 - 11.2
Lunar Transfer 11.5 - 12.0
Mars Transfer 12.5 - 13.5
Interstellar Probe >16.0

The choice o