CATALYST PACK FOR HYDROGEN PEROXIDE

MONOPROPELLANT THRUSTER

KHAFRI IZUAN BIN KHANAFIAH

A thesis submitted in fulfilment of the

requirement for the award of the degree of

Master of Engineering (Mechanical)

Faculty of Mechanical Engineering

Universiti Teknologi Malaysia

FEBRUARI 2014

iii

To my beloved mother and father

iv

ACKNOWLEDGEMENT

In preparing this thesis, I have been contacting with so many people,

researchers, academicians, technicians and other practitioners. All of them have

contributed towards my understanding and thoughts. In particular, I wish to express

my sincere appreciation to my main thesis supervisor, Professor Subramaniam

Krishnan for helping and giving me support, guidance, and critics whenever I needed

in order to finish this under graduate project. I am also thankful to my undergraduate

project seminar panels Professor Wan Khairuddin Wan Ali, Professor Mohammad

Nazri Mohd Jaafar and Professor Shuhaimi Mansor for guiding me towards the

completion of my project with their critics, advices, motivation and support. Without

them, this project will have a lot of overseen flaws.

I am also indebted to a Research Student Grant (RSG) scheme for funding my

Master’s study. Without their support, I will not be able to pursue my study in this

university. I am also indebted to CICT which has provided good internet facilities in

my college. With this facility, I am able to find a lot of information for my research.

My words of appreciation also go to all my friends who have helped me a lot

and lent me their vehicles. I will not forget all of your deeds to me. Last but not least,

I would like to send my gratitude to my beloved father who gave me strength to

pursue this study and gave me guidance along the way when I am here in UTM.

v

ABSTRACT

Silver is one of the common materials used to decompose hydrogen peroxide

but it is heavy and does not last very long while manganese catalyst seems to be

lighter and lasts longer, giving it the potential to replace silver. Unfortunately, this is

yet to be tested. The configuration to produce 100 N thrust by using silver catalyst

needs to be determined, and the feasibility of using manganese calcinated on alumina

needs to be verified. With theoretical calculation and several trial and error

experiments by varying the catalyst pack compaction pressure, injector and nozzle

diameters, catalyst pack heating temperature and propellant tank feeding pressure for

silver catalyst, the configuration to produce 100 N thrust has been obtained. For

manganese catalyst, assorted combinations of alumina sizes and types were tested to

be the catalyst carrier. The method of calcination was also developed in order to get

the highest amount of manganese deposited and it was found that Sasol -alumina

spheres with a diameter of 2.4 mm produce the highest amount of manganese

deposition with an average of 42% after three calcinations using potassium

permanganate as the precursor solution. This was followed by experimental work

which found that the usage of manganese calcinated on the alumina cannot cope with

high pressure in the thruster and tends to break into small pieces and wash out of the

thruster. Silver configuration for producing 100 N thrust has been obtained in this

research. Also, it was found that it is not feasible to use manganese with the method

described in this work.

vi

ABSTRAK

Perak adalah salah satu bahan yang biasa digunakan untuk menguraikan

hidrogen peroksida akan tetapi ianya berat dan tidak kekal lama manakala

pemangkin mangan pula adalah lebih ringan dan tahan lebih lama, memberikan ia

potensi untuk menggantikan perak. Walau bagaimanapun, ini masih belum diuji.

Konfigurasi untuk menghasilkan daya tujah 100 N dengan menggunakan pemangkin

perak perlu ditentukan, dan kesesuaian menggunakan mangan yang dikalsinkan pada

alumina sebagai pengganti perak perlu disahkan. Melalui pengiraan secara teori dan

beberapa ujikaji menggunakan kaedah cuba jaya dengan mengubah tekanan

pemadatan pek pemangkin, diameter lubang penyuntik dan nozel, suhu pemanasan

pek pemangkin dan tekanan suapan tangki bahan pendorong pemangkin perak,

konfigurasi untuk menghasilkan daya tujah 100 N telah diperolehi. Untuk pemankin

mangan pula, pelbagai kombinasi saiz dan jenis alumina diuji untuk menjadi

pengangkut pemangkin. Kaedah pengkalsinan juga telah dibangunkan untuk

mendapatkan jumlah tertinggi mangan yang berjaya dimendapkan dan didapati

bahawa sfera -alumina Sasol dengan garis pusat 2.4 mm menghasilkan jumlah

tertinggi mendapan mangan dengan purata sebanyak 42% selepas tiga kali

pengkalsinan menggunakan kalium permanganat sebagai larutan pendahulu. Ini

diikuti dengan uji kaji yang mendapati bahawa penggunaan mangan yang

dikalsinkan pada alumina ini tidak dapat menampung tekanan tinggi di dalam

pendorong dan cenderung untuk pecah menjadi kepingan kecil dan terkeluar

daripada pendorong itu. Konfigurasi perak untuk menghasilkan daya tujah 100 N

telah berjaya diperolehi dalam kajian ini. Juga, didapati bahawa adalah tidak sesuai

mangan digunakan dengan kaedah yang diperihalkan dalam kajian ini.

vii

TABLE OF CONTENTS

CHAPTER TITLE PAGE

Title page i

Declaration of originality and exclusiveness ii

Dedication iii

Acknowledgements iv

Abstract (English) v

Abstrak (Bahasa Melayu) vi

Table of contents vii

List of tables x

List of figures xii

List of symbols xiv

1 INTRODUCTION 1

1.2 Statement of the Problem 2

1.3 Objective 3

1.4 Scope of Project 3

1.5 Outline of Thesis 3

2 MONOPROPELLANT THRUSTER 5

2.1 History of Rockets 5

2.2 Rocket Fundamentals 6

2.3 Thrust 6

2.4 History of Hydrogen Peroxide Propulsion 11

2.5 Propellant 11

2.6 Monopropellant 12

2.7 Hydrogen Peroxide (H2O2) 13

2.7.1 Structure and Properties 13

viii

2.7.2 Reactions 14

2.7.3 H2O2 Advantages 16

2.7.4 H2O2 Characteristic 17

2.8 Catalyst 19

2.8.1 Types of catalyst 19

2.8.2 Hydrogen peroxide catalyst 20

2.9 Alumina 21

3 METHODOLOGY 23

3.1 Introduction 23

3.2 Experimental Procedure 24

3.3 Thruster 25

3.4 Catalyst Pack Preparation 25

3.4.1 Silver Catalyst 26

3.4.2 Ceramic Catalyst 27

3.4.3 Test Facility 28

3.5 Rocket Performance Calculation 32

3.5.1 Corrected Pressure Plot 32

3.5.2 Estimation of Vacuum Specific Impulse 34

3.5.3 Estimation of Vacuum Thrust 36

3.5.4 Performance Calculation after Burnout 38

3.5.5 Catalyst Pack Pressure Drop and Mass

Flux Calculation

41

3.6 Testing 43

3.6.1 Operating Test Rig Procedure 43

3.7 Types of Testing 45

3.7.1 Nitrogen Testing 45

3.7.2 Nitrogen and Water Testing 45

4 THRUSTER DESIGN AND FABRICATION 47

4.1 Introduction 47

4.2 Thruster Design and Fabrication 47

4.2.1 Design Calculation 50

ix

4.2.2 Program 53

4.2.3 Output 54

4.3 Drawings 56

4.4 Fabrication 67

4.5 Finished Product 68

5 RESULTS AND DISCUSSION 71

5 .1 Test Runs for Silver catalyst 71

5.2 Calcination Trials for Ceramic catalyst 79

5.3 Test Runs for Ceramic catalyst 80

6 CONCLUSION AND RECOMMENDATION 83

6.1 Conclusions 83

6.2 100N thruster using silver catalyst 83

6.3 Ceramic based catalyst 84

6.4 Recommendations 85

REFERENCES 86

APPENDIX 91

Appendix A – Fortran Programming 91

Appendix B – Hot Test Results of 100N H2O2

Monopropellant Thruster Adopting Silver

Catalyst Pack

95

Appendix C – Sasol Alumina MSDS 107

x

LIST OF TABLES

TABLE NO. TITLE PAGE

2.1 Comparison of oxidation potential of different oxidant 16

2.2 Uses of Alumina 22

3.1 The output values of the program after calculation by

using FORTRAN programming

24

3.2 Designed and actual dimensions of nozzle diameter

together with calculated nozzle throat area ratio

24

3.3 Procedure of preparing calcinated ceramic catalyst

using potassium permanganate as a precursor

27

3.4 Types of ceramic samples with its manufacturer 28

4.1 The output values of the program after calculation by

using FORTRAN Programming

55

4.2 Summary of calculated crucial thruster dimensions 56

5.1 Summary of test runs 71

5.2 Summary of test results for series I using H2O2 of

concentration of 0.88, injector orifice diameter of

0.914mm, and nozzle throat diameter of.10.007

72

5.3 Summary of test results for series II using H2O2 of

concentration of 0.90, injector orifice diameter of

1.215mm, and nozzle throat diameter of 10.007

72

5.4 Summary of test results for series III using H2O2 of

concentration of 0.90, injector orifice diameter of

1.898mm, and nozzle throat diameter of 6.789

73

5.5 Summary of calcination trial on several types of

alumina

79

xi

5.6 Summary of all the tests using ceramic based catalysts 81

5.7 Summary of successful tests using catalyst pack

containing Sasol alumina pellets

82

xii

LIST OF FIGURES

FIGURE

NO.

TITLE PAGE

2.1 Pressure balance on chamber and nozzle interior walls 7

2.2 Molecular Structure of H2O2 in gas and solid state 14

3.1 Circular mesh of 20 x 20 silver mesh 26

3.2 Rocket thruster test facility 29

3.3 Schematic diagram of rocket thruster test facility 30

3.4 Control panel that will be used to control the test facilities

at a safe distance

31

3.5 Typical record of test through LABView software 32

3.6 Derivatives of injection pressure and nozzle entry pressure 33

3.7 Plot of corrected absolute pressure 34

3.8 Example of a complete profile of thrust and mass flow rate

from start until burnout

41

3.9 Example of mass flux and pressure drop across catalyst

pack versus time

42

3.10 Example of mass flux versus pressure drop across catalyst

pack graph

43

3.11 Summary of methodology 46

4.1 100N thruster assembly drawing 57

4.2 Cross section of the 100N thruster with labels drawing. 58

4.3 100N thruster injector with cross section drawing 59

4.4 100N thruster body details drawing 60

4.5 100N thruster nozzle details drawing 61

4.6 100N thruster nozzle holder details drawing 62

4.7 Injector distribution plate details drawing 63

4.8 Nozzle distribution plate details drawing 64

xiii

4.9 100N thruster body with 105mm combustion chamber

details drawing

65

4.10 100N thruster body with 155mm combustion chamber

details drawing

66

4.11 An illustration of a 100N thruster assembly 68

4.12 Complete 100N thruster after fabrication with labels 68

4.13 Finish product after fabrication 69

5.1 Time differentials of injection pressure dtdP2 and nozzle

entry pressure dtdP4 versus time

75

5.2 Corrected pressures versus time 75

5.3 Calculated thrusts and propellant flow rates versus time 76

5.4 Pressure drop across catalyst pack for two compaction

pressures cpP .

76

5.5 Variation of

*

c with respect to that i

incp

d

PT

.

78

5.6 Variation of ignition delay with respect to

i

incp

d

PT

.

78

5.7 Time differentials of injection pressure and nozzle entry

pressure versus time for first test, Table 5.7

82

5.8 Measured pressure traces versus time for the test 1, Table

5.7.

82

5.9 Calculated thrusts and mass flow rates versus time for test 1

from Table 5.7.

82

5.10 Time differentials of injection pressure and nozzle entry

pressure versus time for successful test 2 from Table 5.7

82

5.11 Measured pressure traces of test 2 from Table 5.7 82

5.12 Calculated thrusts and mass flow rates versus time for test

2, from Table 5.7

82

xiv

LIST OF SYMBOLS

CPA = Cross sectional area of catalyst pack

eA = Nozzle exit area

dc = Coefficient of discharge for the injector orifice

𝑐∗𝑒𝑥𝑝 = Experimental characteristic velocity

𝑐∗𝑡ℎ𝑒𝑜 = Theoretical characteristic velocity

0FC = Characteristic thrust coefficient

expFsl

C

= Experimental sea level thrust coefficient

theoslFC

= Theoretical sea level thrust coefficient

id = Injector orifice diameter

CPD = Catalyst pack diameter

eD = Nozzle exit plane diameter

tD = Nozzle throat diameter

F = Thrust

expslspI = Sea level specific impulse

CPL = Catalyst pack length

m = Molar mass of decomposed gas

xv

pm = Propellant mass flow rate

ap = Ambient pressure at sea level

ep = Nozzle exit plane pressure

CPp0 = Catalyst pack entry pressure

ip0 = Injection pressure

𝑃0𝑛 = Nozzle entry stagnation pressure

fhpp0 = High pressure tank final pressure

ihpp0 = High pressure tank initial pressure

ptp0 = Propellant tank pressure

uR = Universal gas constant, 8314.3 j/kmol-K

adT0 = Adiabatic flame temperature

nT0 = Nozzle entry stagnation temperature

hptV = High pressure tank volume

ptV = Propellant tank volume

= Ratio of specific heats for the neutral gas in the high pressure tank

= Nozzle area ratio, te AA

CPp = Catalyst pack pressure drop

ip = Injector pressure drop

oop = On/off valve pressure drop

PRp = Minimum pressure difference across pressure regulator

xvi

Ft = Thrusting time

rtt = Residence time in catalyst pack

η𝑐∗ = *c efficiency

η𝐶𝐹 = FC efficiency

n0 = Nozzle entry stagnation density

p = Propellant density

= Gas mass flux through the catalyst pack

CHAPTER 1

INTRODUCTION

Liquid rocket propellant system commonly divided into two types which are

monopropellant and bipropellant [1]. In a monopropellant rocket system, a single

liquid system is used as propellant. The most commonly used monopropellant is

hydrazine (N2H4) which is highly toxic and very unstable unless handled in solution

[1]. While in bipropellant rockets, two types of liquid system are used. The liquids

are categorized into two which are the fuel and the oxidizer. The most common

combination of bipropellant are monomethylhydrazine (MMH) and nitrogentetroxide

(N2O4) both of which are highly toxic and unstable [1-2]

The above propellant mixtures require particular propellant handling and

prelaunch preparation. Because of the requirement, hydrazines and nitrogen tetroxide

have become less attractive fluid while hydrogen peroxide (H2O2) seems better as

potential alternative [3-7]. Wernimont E. J. [8] stated that the non-toxic chemical

which is rocket grade H2O2 (concentration greater than 85%) has a natural familiarity

to human chemistry thus it is the best wide-ranging solution for space, land, air and

sea applications. When choosing H2O2 as a monopropellant, the other benefits are the

significant cost saving and simplification of health and safety precautions needed

throughout the fabrication, storage and handling of the propellants [2].

The cost of manufacturing and preparing the thruster does not scale down

proportionally with the thruster size due to the advantages. These advantages have

2

special relevance to low or medium thruster. The propulsion mechanism was derived

from silver meshes that act as a catalyst for the decomposition of H2O2. Since

decomposition products are safe, H2O2 monopropellant also can be considered as

safe. The reaction equation for the decomposition process is as per equation (1.1):

HeatgOlOHlOH Catalyst )()(2)(2 2222 1.1

Equation (1.1) shows that superheated steam and oxygen with heat are

released from the decomposition process. It means that no other lethal gas is released

to the atmosphere. Based on this fact, concentration up to 90% of the H2O2 rocket

grade was prepared. In order to achieve higher specific impulse which can give more

thrust, higher concentrations of H2O2 were needed. Unfortunately, silver catalyst

cannot withstand the heat generated by the decomposition process while adopting

high concentration of H2O2. By introducing alumina coated with manganese oxide,

experiments were conducted in order to validate this catalyst for satellite propulsion

usage. These ceramic based catalysts are expected to be lighter, cheaper, and of

longer life.

1.2 Statement of the problem

For the improved application of hydrogen peroxide in rocket propulsion,

major research activities are presently moving towards indentifying suitable catalyst

system to decompose hydrogen peroxide. Traditionally screen of pure silver or silver

coated nickel or stainless steel have been adopted. Silver based catalysts are heavy

and of short life. Furthermore, when the concentration of hydrogen peroxide

increases (to realize higher specific impulse), the silver based system cannot

withstand high temperatures (>1000K) of the decomposed products of hydrogen

peroxide. In searching for suitable alternative catalyst, manganese oxide calcinated

on the substrates of alumina, titanium oxide, or cordierite has been shortlisted as the

candidates for the catalyst system. These ceramic based catalysts are expected to be

3

lighter, cheaper, and of longer life. However, detailed characterization of the system

has not been reported.

1.3 Objective

1. To determine the sustainability of the conventional silver catalyst in the

development of a H2O2 monopropellant rocket engine of 100N thrust.

2. To find out the feasibility of using alumina based catalyst in the development

of monopropellant thruster.

1.4 Scope of Project

In order to achieve the objectives of the project, several scopes have been

adopted. The scopes include using FORTRAN programming to ease the calculation

of the parameters needed for the design, designing and fabricating the thruster using

the parameters that have been calculated. It also include preparing silver and ceramic

catalyst using methods applied by others and developing own method of preparation

for ceramic catalyst.

1.5 Outline of Thesis

This thesis consists of six chapters. In this chapter, introduction, objective,

statement of problems, scope of this project and summary of works are reviewed.

While in Chapter 2, theory and literature reviews on hydrogen peroxide and its

catalyst, thruster, micro thruster and its applications from various resources are

summarized.

In Chapter 3, the discussion is on the methods of completing this project

using hardware and software implementation together with catalyst preparation and

4

testing procedures. The design process is elaborated in Chapter 4. The entire

experimental results and discussion are described in Chapter 5. Lastly, Chapter 6,

encompasses conclusion of this project and future work which can be done hence

recommended.

86

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