Design and Prototyping of a Solid Fuel/Liquid Oxidizer
Hybrid Rocket Engine
By
Hammad Ali Shah(ME122011)
Issac Akram (ME122013)
Shahid Khizar (ME122042)
Sidra Jabeen Siddique (ME122045)
DHA Suffa University
Karachi, Pakistan
© 2016, Hammad Ali Shah, Dr. Bilal A. Siddiqui
1 Abstract | Team ISSH
CERTIFICATE
Submitted in partial fulfillment of the requirement of degree of Bachelor of
Engineering (Mechanical Engineering)
Design and Prototyping of a Solid Fuel/Liquid Oxidizer Hybrid
Rocket Engine
BATCH – 2012
Name
Registration No.
Hammad Ali Shah (Project Leader)
Issac Akram
Shahid Khizar
ME-122011
ME-122013
ME-122042
Sidra Jabeen Siddique
ME-122045.
Project Supervisor
Dr.Bilal Siddique
Assistant Professor
Primary Examiner
Dr. Johar Farooqi
Professor and Dean (EAS)
FACULTY OF ENGINEERING AND ACADEMIC SCIENCES
DHA SUFFA UNIVERSITY
2016
2 Abstract | Team ISSH
This page was left intentionally
3 Abstract | Team ISSH
Abstract
Propulsion is generally a mature science, but one technology has recently generated a lot
of research and development by offering safer and cheaper means of propulsion. It is called
the hybrid rocket technology, which has transformed the space industry during the last two
decades. Unfortunately, this technology has not been given its due share in academia and
industry outside the Western hemisphere.
This work describes the development of a hybrid rocket propulsion system, as part of the
first phase of DSU’s pioneering hybrid rocket engine (HRE) program, the first of its kind
in Pakistan and the third in Asia. The research objective is to design and develop
methodologies for design, development and testing of hybrid rocket technology.
We present the theoretical tools for the development of the design algorithm of HREs,
implemented in MATLAB programming language. Based on these calculated parameters,
rocket engine drawings are developed, taking into account the practical aspects of
indigenous manufacturing. Finally, the manufacturing scheme, testing procedures, data
acquisition, fuel grain processing, test results and post-test analysis are presented. We
conclude with recommendations for future research and improvements in the current
scheme.
4 Abstract | Team ISSH
Acknowledgement
First and foremost, our appreciation is addressed to our supervisor, Dr. Bilal Ahmed Siddique
for his valuable practical knowledge, advices, discussions and support that made this work
possible.
We are also thankful to our co-supervisors, Engr. Shoib Ahmed and Dr. Rida Ahmed who
shared stimulating ideas with us.
We would also like to thank our collaborators at HEJ Center of Chemical and Biological
Sciences, University of Karachi for chemical characterization of solid fuel used in our
calculations. Finally, Mr. Obaid Rahmani, Mr. Anas Iftikhar and Mr. Shan Ahmed of AeroX
must be thanked for their help in giving us tips about manufacturing aspects of the project.
We would like to thank our Dean, Dr. Johar K. Farooqi and Vice Chancellor, Dr. Sarfraz
Hussain for invaluable support and unfailing encouragement.
We are also thankful to Mr. Mudassir Qadeer and other faculty members who let us use the
experimental facilities and labs in DHA Suffa University and by helping us out throughout
the project by sharing their information with us. Our juniors Khizar, Umer, Waqas and Mohsin
are working on a similar project, and helped us with testing and provided the igniter, for which
we are indebted.
Finally, our special thanks to our families for their financial and motivational support and
encouragement they gave us which made it possible for us to progress in our work.
5 Abstract | Team ISSH
Table of Contents
Abstract .............................................................................................................................. 4
Table of Contents .............................................................................................................. 6
List of Tables ..................................................................................................................... 8
List of Figures .................................................................................................................... 9
Nomenclature .................................................................................................................. 12
Chapter 1:
Introduction ........................................................................................... 15
1.1
Motivation .......................................................................................................... 15
1.2
Objective ............................................................................................................ 17
1.3
Statement of Requirements ................................................................................ 18
1.4
Project Timeline ................................................................................................. 27
Chapter 2:
Literature Review .................................................................................. 29
Chapter 3:
Design Methodology .............................................................................. 34
3.1
Design Process Flow .......................................................................................... 34
3.2
Design Parameters .............................................................................................. 34
3.3
HRE Design Software ........................................................................................ 42
3.4
Mechanical Design of Rocket Components ....................................................... 43
Chapter 4:
4.1
Manufacturing of Designed Rocket ..................................................... 53
Manufacturing and Instrumentation Costs ......................................................... 53
6 Table of Contents | Team ISSH
4.2
Manufacturing Details ........................................................................................ 54
4.3
Fuel Grain Casting ............................................................................................. 67
Chapter 5:
User Manual with Designed HRE Shaheen ........................................ 72
5.1
Engine Specifications ......................................................................................... 72
5.2
Package Part List ...................................................................................... 73
5.3
Assembling Procedure........................................................................................ 75
5.4
Post Firing Cleaning Procedures ............................................................... 78
5.5
Safety and First Aid............................................................................................ 79
5.6
Disclaimer .......................................................................................................... 79
Chapter 6:
Experimental Results and Discussion .................................................. 80
6.1
Test Procedure .................................................................................................... 80
6.2
First Static Test................................................................................................... 83
6.3
Second Static Test .............................................................................................. 87
6.4
Third Static Test ................................................................................................. 90
6.5
Testing Videos.................................................................................................... 95
Chapter 7:
Design Recommendations ..................................................................... 96
Chapter 8:
Conclusion and Future Work .............................................................. 97
References ........................................................................................................................ 98
7 Table of Contents | Team ISSH
List of Tables
Table 1 Statement of Project Requirements ............................................................................................. 18
Table 2 Landmark researches in Hybrid Rocket Technologies .............................................................. 30
Table 3 Costs of Manufacturing ................................................................................................................. 53
Table 4 Engine Specification ...................................................................................................................... 72
Table 5 Test Fire Matrix ............................................................................................................................. 80
8 List of Tables | Team ISSH
List of Figures
Figure 1 Schematic of a Hybrid Rocket Engine ........................................................................................ 16
Figure 2 Gantt Chart of Project Deliverables ........................................................................................... 28
Figure 3 Flow Chart of Design Process ..................................................................................................... 35
Figure 4 HRE Design Program, code sample. Copyrights Dr. Bilal 2016 .............................................. 42
Figure 5 Sample Output of HRE Design Program, Copyrights Dr. Bilal 2016 ...................................... 43
Figure 6 Combustion Chamber (all dimensions in mm) ......................................................................... 44
Figure 7: Final Combustion Chamber....................................................................................................... 45
Figure 8 Injector Bottom Retainer (all dimensions in mm) ..................................................................... 46
Figure 9: Final Injector Bottom Retainer ................................................................................................. 47
Figure 10 Injector Top Retainer ................................................................................................................ 48
Figure 11: Final Injector Top Retainer ..................................................................................................... 48
Figure 12 Injector (all dimensions in mm) ................................................................................................ 49
Figure 13 Final Injector .............................................................................................................................. 50
Figure 14 Convergent Nozzle (all dimensions in mm) .............................................................................. 50
Figure 15: Nozzle ......................................................................................................................................... 51
Figure 16 Nozzle Retainer (all dimensions in mm) ................................................................................... 52
Figure 17 Final Nozzle Retainer ................................................................................................................. 52
Figure 18 Combustion Chamber design (all dimensions in mm) ............................................................ 54
Figure 19 Manufactured Combustion Chamber ...................................................................................... 56
Figure 20 Nozzle design (all dimensions in mm) ....................................................................................... 57
Figure 21 Manufactured Nozzle ................................................................................................................. 58
Figure 22 Injector design (all dimensions in mm) .................................................................................... 59
Figure 23 Manufactured Injector .............................................................................................................. 60
Figure 24 Injector Top Retainer design (all dimensions in mm, degrees) ............................................. 61
Figure 25 Manufactured Injector Top Retainer ....................................................................................... 63
Figure 26 Nozzle Retainer design (all dimensions in mm) ...................................................................... 64
9 List of Figures | Team ISSH
Figure 27 Manufactured Injector Bottom Retainer ................................................................................. 65
Figure 28 Assembled top and bottom of retainer with bolts .................................................................... 65
Figure 29 Nozzle Retainer design (all dimensions in mm) ....................................................................... 66
Figure 30 Manufactured Nozzle Retainer ................................................................................................. 67
Figure 31 Pentacosane packed in bags with rubber sheet and other apparatus used for casting ........ 68
Figure 32 Casting process in progress ....................................................................................................... 69
Figure 33 Molten wax being poured into the combustion chamber through funnel ............................. 69
Figure 34 Combustion chamber view after casting .................................................................................. 70
Figure 35 Rubber sheet acting as insulator ............................................................................................... 70
Figure 36 Fuel grain fitted inside the combustion chamber .................................................................... 71
Figure 37 Shaheen Hybrid Rocket Engine Schematic .............................................................................. 72
Figure 38 Oxidizer Tank ............................................................................................................................. 74
Figure 39 Fuel Grain ................................................................................................................................... 74
Figure 40 Condi Nozzle ............................................................................................................................... 75
Figure 41 Nozzle Divergent portion ........................................................................................................... 76
Figure 42 Combustion Chamber ................................................................................................................ 76
Figure 43 FrSky Taranis X9D Plus Transmitter ...................................................................................... 78
Figure 44 Oxidizer tank with needle valve regulator ............................................................................... 83
Figure 45 First static test in progress ........................................................................................................ 84
Figure 46 Injector Plate (back view, post test) .......................................................................................... 85
Figure 47 Injector Plate (front view, post test) ......................................................................................... 85
Figure 48 Injector Top Retainer (post test) ............................................................................................... 85
Figure 49 Fuel Grain Single Section after 1st test ..................................................................................... 86
Figure 50 Complete Fuel Grain section view after 1 st test ....................................................................... 87
Figure 51 X-Ray of Fuel Grain (post test) ................................................................................................. 88
Figure 52 Visual Inspection of Test Stand (post test) ............................................................................... 89
Figure 53 Test being carried out in the test pit ......................................................................................... 89
Figure 54 CAD model of car chassis mounted with oxidizer tank and engine ....................................... 90
10 List of Figures | Team ISSH
Figure 55 1000 psi pressured gas can be clearly seen on the gauge ........................................................ 90
Figure 56 3rd test being carried out in the test pit ................................................................................... 91
Figure 57 Weighing machine fitted with Salter Precision Machining Inc load cells ............................. 91
Figure 58 Nozzle and its Retainer (after third test) .................................................................................. 92
Figure 59 Graphite Nozzle Insert with Residue Deposits ........................................................................ 93
Figure 60 Inside View of Combustion Chamber with Wax and Rubber Residue ................................. 93
Figure 61 Nozzle Retainer after Third Test .............................................................................................. 94
Figure 62 Residue (unburnt wax) after 3rd test ....................................................................................... 94
11 List of Figures | Team ISSH
Nomenclature
A
Area
[m2]
c*
Characteristic velocity
[m/s]
Cd,inj Injector discharge coefficient [dimensionless]
Cd,nozz Nozzle discharge coefficient
[dimensionless]
d
Diameter
[m]
dinj
Injector plate diameter
[m]
F
Thrust
[N]
go
Gravitational acceleration
[m/s2]
Isp
Specific impulse
[s]
Lp
Length of port
[m]
L
Length
[m]
M
Molecular weight
[kg/mol]
m
Mass
[kg]
ṁ
Mass flow rate
[kg/s]
n
Number of holes
[dimensionless]
OF
Oxidizer fuel ratio
[dimensionless]
12 Nomenclature | Team ISSH
P
Pressure
[N/m2]
RA
Universal gas constant
[J/kmol.°K]
Sy
Yield strength
[MPa]
Sf
Factor of safety
[dimensionless]
T
Temperature
[K]
tw
Wall thickness
[m]
tb
Burn time
[s]
Ve
Exit velocity
[m/s]
Greek Symbols
ρ
Density
[kg/m3]
ε
Nozzle expansion ratio
[dimensionless]
β
Nozzle convergent angle
[degrees]
α
Nozzle divergent angle
[degrees]
Specific heat ratio
[dimensionless]
Volume
[m3]
∀
13 Nomenclature | Team ISSH
Subscripts
c
combustion chamber
con
nozzle convergent section
d
design
div
nozzle divergent section
e
exit exhaust
f
fuel
inj
injector
i
internal
ox
oxygen
p
port
t
nozzle throat
w
nozzle wall
Acronyms
HRE
hybrid rocket engine
LOX liquid oxygen
GOX gaseous oxygen
14 Nomenclature | Team ISSH
Chapter 1:
Introduction
1.1 Motivation
Rockets afford the technological means of transporting man and material over long
distances and great altitudes. Rocket science is the pinnacle of engineering in terms of the
challenges of interdisciplinary considerations during design and parameters being on the
extreme limit of nature. Therefore, mastering it is no simple feat.
For combustion to occur, at least two components are required: a source of hydrocarbons
(fuel) and a source of oxygen (oxidizer). Until recently, there have been only two types of
successfully deployed rocket systems: solid rocket motors (SRMs) and liquid rocket
engines (LREs). Both fuel and oxidizer are in solid form in SRMs, whereas they in liquid
form in LREs. Since around 2000 AD, there has been a keen interested in a previously
inefficient type of rocket called the hybrid rocket engine (HRE) where one of the
components (usually fuel) is in solid form, whereas the other component (usually oxidizer)
is either in liquid or gaseous form.
HREs have combined the best qualities of both SRMs and LREs. Some of these are:
1) Low cost, much lower than SRMs.
2) Thrust to weight flow ratio (Isp) higher than SRMs, but lower than LREs.
3) Ability to be throttled.
4) Safety, as opposed to the often-explosive nature of SRMs.
5) Storability, if there are no cryogenic components.
15 Introduction | Team ISSH
6) Availability of propellants.
7) Regression rates now comparable to SRMs.
8) Very less complicated plumbing as opposed to LREs.
9) Ability to restart.
10) Reusability.
Figure 1 Schematic of a Hybrid Rocket Engine
With so many advantages, HREs have enthralled the scientific community across
disciplines for more than a decade now. The technology has been demonstrated for
suborbital space flight and is being considered for satellite launch as well. Pakistan is
actively pursuing its space program, and needs cheap alternative to conventional rocket
launch in order to commercialize its space. HREs is the answer to this quest. Hence, this
thesis aims to break the ice on this subject in the country.
16 Introduction | Team ISSH
1.2 Objective
The main objectives of the thesis include the following aspects:
1. Developing an algorithm for design of salient parameters of HREs.
2. Developing a computer code to optimize HRE design based on the algorithm in
Objective 1 above.
3. Develop manufacturing drawings for the optimum design.
4. Manufacture the designed rocket.
5. Verify and further optimize based on static testing of rocket design.
17 Introduction | Team ISSH
1.3 Statement of Requirements
The following SOR was approved for this project. Abbreviations D, W, L, H and I
stand for desired, wish, low, high and impossible. This requirement is a fundamental
criterion for judging project outcomes.
Table 1 Statement of Project Requirements
Title: Design and Prototyping of
a Solid Fuel/Liquid Oxidizer
Hybrid Rocket Engine
CHANGES
D/W
REF
Issue: 01
Date: Day-April-2016
REQUIREMENTS
1
Introduction
1.1
Preamble
A novel propulsion technology called Hybrid Rocket Engine
(HRE) uses the fuel and oxidizer in two different phases
(mostly solid fuel and liquid oxidizer) to generate thrust,
while conventional technology uses both the fuel and
oxidizer in the same phase (solid in the case of Solid Rocket
(SRE) and liquid in the case of Liquid Rocket Engines
(LRE).
This new technology has the potential to have a low carbon
footprint and have the advantages of both solid and liquid
rocket engines. This field has attracted a lot of research in
18 Introduction | Team ISSH
universities and industry lately. SpaceShipX programs which
were manufactured with intention of space exploration use
Hybrid Rocket Engine. Bloodhound SSC will also be
equipped with a rocket engine which would help it to reach
1,000 mph on land. Various manufacturers like Space
Propulsion Group, Scaled Composites, Space Dev/ Sierra
Nevada Corp have already succeeded in bringing the concept
to fruition.
This study aims at designing and manufacturing a complete
hybrid rocket engine. This will be a simple, cheap and safe
hybrid engine at the lab scale.
1.2
Scope
Hybrid Rocket Engine is a safe and simple solution for space
exploration and high speed ground travel.
1.3
Related Documents
1.3.1
Books
Space Mission Analysis and Design. 3rd Edition, Edited by
Wiley J. Larson and James R. Wertz
19 Introduction | Team ISSH
Rocket Propulsion Elements, 7th Edition, George P. Sutton
and Oscar Bublarz
Fundaments of Hybrid Rocket Combustion and Propulsion,
By Martin J. Chiaverini and Kenneth K. Kuo
1.3.2
Software
1.3.2.1
Solid Works
1.3.2.2
Ansys 13.0
1.3.2.3
Rocket Sim
1.3.3
Research Papers
Preliminary Study of a Hybrid Rocket Pedro Paulo de
1.3.3.1
Oliveira Alcaria Guerreiro
Hybrid Rocket Motor by Zach Arena, Alexander Athougies,
1.3.3.2
and Alden Rodulfo
High Performance Hybrid Upper Stage Motor Arif
Karabeyoglu, Jose Stevens, Dmitriy Geyzel, Brian Cantwell§
1.3.3.3
Space Propulsion Group Inc., Sunnyvale CA and Dave
Micheletti MADA, Butte MT
20 Introduction | Team ISSH
1.4
Definitions, Abbreviations and Symbols
1.4.1
Definitions
1.4.1.1
Thrust rise time
Time taken by system to go from zero thrust to full thrust.
1.4.1.2
Specific Impulse (Isp)
It is impulse delivered per unit weight of propellant
consumed
1.4.1.3
Total impulse (I)
The change in momentum that can be accomplished by the
motor, expressed in ‘Newton-second' (Ns).
1.4.1.4
Chamber pressure (Pc)
The pressure in the combustion or reaction chamber.
1.4.2
Abbreviations
LOX=Liquid Oxygen
GOX=Gaseous Oxygen
HTPB=Hydroxyl Terminated-Polybutadiene
APCP=Ammonium Perchlorate Composite Propellant
21 Introduction | Team ISSH
1.4.3
Symbols
D
Demand
(A mandatory requirement)
W(H) Wish high
(A highly desirable attribute)
W(L) Wish low
(A low desirable attribute)
W(I)
Wish Impossible(An attribute impossible without
funding)
2
Technical Requirements
2.1.1
Good engineering practices
2.1.2
Understanding and use of given software
2.1.3
Lab space for engine testing
2.1.4
Knowledge
of
Gas
Dynamics,
Mechanics
of
Materials,
Organic
Materials
Chemistry,
Engineering,
Thermodynamics, Measurements and Instrumentation
2.1
Description & Purpose
To design and develop a safe, cheap, reliable rocket engine
with low carbon footprint
2.1.1
Engine design will cater for thrust termination and restart.
2.2
Functional Characteristics
22 Introduction | Team ISSH
Manufacture and statically test the designed hybrid rocket
W(L)
engine at laboratory scale
Design and analyze a hybrid rocket engine with thrust of 400
D
2.2.1
N
W(H)
2.2.2
Use paraffin wax as fuel and liquid oxygen as oxidizer
W(H)
2.2.3
Generate a thrust of nearly 1000N
W(I)
2.2.4.1
Test the rocket in flight
W(H)
2.2.4.2
Launch our rocket with payload
Test multiple engines with Ammonium Perchlorate,
W(H)
2.2.5.1
HTPB/NO2, and sugar based propellant and compare their
performance with hybrid rocket designed.
Develop a test stand with appropriate instrumentation for
W(L)
2.2.5.2
static testing
2.3
Physical and other Characteristics
All rocket components should be visible and not covered
W(H)
2.3.1
from sight or reach due to test stand attachments.
D
2.3.2
Rocket design with fail-safe features and safety alarms.
W(H)
2.3.3
All connections should be tight and leak proof.
23 Introduction | Team ISSH
2.4
Design & Construction
2.4.1
The main components of Hybrid Rocket Engine are

Fuel Grain

Oxidizer Tank

De Laval Nozzle

Orifice Plate

Injector

Pressure Vessel
W(L)
2.4.2
Above components used are locally available.
W(L)
2.4.3
LOX and paraffin are readily available.
W(H)
2.4.4
Graphite disks for machining nozzle inserts are available.
Ammonium
W(I)
perchlorate’s
availability for
2.2.5.1
is
2.4.5
questionable
2.5
Environmental Conditions
The rocket motor can function under standard environmental
W(L)
2.5.1
conditions without the level of pollution caused by SREs and
LREs
24 Introduction | Team ISSH
2.6
Reliability & Maintenance
The fuel grain needs to be replaced every time after use and
W(H)
2.6.1
the motor should be checked for structural integrity.
2.7
Safety
Fire-fighting equipment should be available at the test site
W(H)
2.7.1
with trained fire fighters.
W(H)
2.7.2
Testing should be done with minimal public attendance.
3
Quality
The project process will be monitored by means of weekly
3.1
project meetings.
3.2
The project plan will be monitored by means of a Gantt chart.
The chosen design will be presented at the concept design
3.3
board.
The chosen design will be assessed by rating and weighing
3.4
matrices.
The technical details will be scrutinized before detailed
3.5
design work is undertaken by means of the design review
board.
25 Introduction | Team ISSH
4
Miscellaneous
The Project Content will be monitored by the project
4.1
advisors.
Hammad Ali Shah has been appointed as the project leader
4.2
and is to ensure that work proceeds according to the project
plan.
5
Costs
The estimated cost of the Project is Rs. 100,000/= (0.1
5.1
Million)
The above mentioned cost is subjected to material cost,
5.2
assembly requirements, testing, calibration; cost may vary
due to unavailability of some equipment or materials
The project has been submitted for sponsorship from Space
5.3
and Upper Atmosphere Research Commission (SUPARCO)
of Pakistan
In case of unavailability of fund the W(I) won’t be
5.4
completed
6
Constraints
6.1
Cost and funding are the major constraints of the project.
26 Introduction | Team ISSH
6.2
Permission for testing has been sought from SUPARCO.
6.3
For oxidizer like NO2 we might face availability issues
Engr. Dr.
Bilal
Project Advisors/Co-Advisors
Engr. Shoib
Ahmed
Siddiqui
(Advisor)
Dr.
Rida
Ahmed
Advisor)
(Co-Advisor)
1.4 Project Timeline
The timeline of deliverables in the project is shown in the Gantt Chart of Figure 2.
27 Introduction | Team ISSH
(Co-
Figure 2 Gantt Chart of Project Deliverables
28 Introduction | Team ISSH
Chapter 2:
Literature Review
The United States Air Force (USAF) developed different hybrid rocket propulsion systems
for sounding rockets and tactical missiles. In January 1994, the United States Air Force
Academy successfully launched a 6.4-m (21 feet) long, 2670-N (600-lbf) thrust
LOX/HTPB hybrid rocket to reach 3-km (10000 feet) altitude. In the next year, this rocket
was developed to deliver 3570-N (800-lbf) for 15 s to reach 4.6 km (15000-feet) altitude.
In 1997 the Academy developed an 890-N (200-lbf)-thrust hybrid-rocket propulsionsystem that used N2O and polyethylene. The Academy studied the auto ignition in hybrid
rocket that used H2O2 and the regenerative cooling of hybrid rocket nozzle using N2O.
Alkuam and Alobaidi [7] recently produced a very comprehensive review of experimental
and theoretical research in hybrid rocket engine technology. This section is mostly
summarized from that survey. Table 2 lists the landmark studies in this technology.
Even though the basic experiments in hybrid rocket engines had started somewhere around
1930-1950, but it was quickly abandoned due to low thrust and regression rates. However,
this changed at the turn of the century (1990-present), when Dr. Arif Karabeyoglu at
Standford discovered very favorable burn characteristics of paraffin wax. In 1999, NASA
has awarded a contract to Lockheed Martin Astronautics, along with subcontractors Boeing
Rocketdyne and Thiokol, to begin development of a peroxide-oxidized hybrid motor for
upper stage application in reusable launch vehicles [18]. This led to resurgence in interest
in this technology. Hudson et al [8] used a 2”x10” lab scale hybrid rocket motor for
spectroscopic measurement of rocket plumes [14].
29 Literature Review | Team ISSH
Table 2 Landmark researches in Hybrid Rocket Technologies
1995 D. J.
Optimization of Hybrid-Rocket-Booster Fuel-Grain Design
Vonderwell
et al.
1998 P. L. Vergez
Tactical Missile Guidance with Passive Seekers Under High
Off-Boresight Launch Conditions
1998 M. K.
Hudson et al.
2000 A. M. Wright
et al.
2000 R. Shanks et
UV, visible, and infrared spectral emissions in hybrid rocket
plumes
The effect of high concentration guanidinium azo tetrazolate on
thrust and specific impulse of a hybrid rocket
Hybrid Rocket Motor for Instrumentation Studies
al. A
Labscale
2001 M. F.
A ground test rocket thrust measurement system
Desrochers et
al.
2003 M. A.
Karabeyoglu
Development of High-Burning-Rate Hybrid-Rocket-Fuel Flight
Demonstrators
et al.
2003 A.
Karabeyoglu
Scale-up tests of high regression rate liquefying hybrid rocket
fuels
et al.
30 Literature Review | Team ISSH
2004 M. K.
Hudson et al.
2004 G. A. Risha
et al.
2005 A.
Karabeyoglu
Guanidinium Azo-Tetrazolate (GAT) as a High Performance
Hybrid Rocket Fuel Additive
Characterization of nano-sized particles for propulsion
applications
Design of an Orbital Hybrid Rocket Vehicle Launched from
Canberra Air Platform
et al.
2011 D. B. Larson
et al.
Characterization of the Performance of Paraffin/LiAlH4 Solid
Fuels in a Hybrid Rocket System
2013 A. Sossi et al. Combustion of HTPB-Based Solid Fuels Loaded with Coated
Nano-aluminum
Guanidinium Azo-Tetrazolate (GAT) was added to hydroxyl-terminated polybutadiene
(HTPB) in two concentrations by mass: 15% and 25%. Although the thrust achieved was
increased with the addition of GAT, and both concentrations created roughly equal
increases in thrust, the specific impulse was lowered somewhat with the additive present
[13]. Due to the presence of positively and negatively charged chemical components, the
GAT salt produced faster regression when added to the HTPB. Nano-aluminum (nAl)
particles were characterized for their effect on the regression rates of the solid fuel
component of hybrid rockets. Particle sizes were 50 nm and 100 nm, coated with a number
of different shielding organic reagents, then added to “hydroxyl-terminated polybutadiene
(HTPB)-based solid fuels” [17] (A. Sossi et al. ). The resulting combustion data were
analyzed to generate a “continuous time-resolved regression rate” [17]. The coated nAl
31 Literature Review | Team ISSH
particles were found to increase the performance of nAl-enhanced HTPB preparations
when compared to unalloyed HTPB when burned in the presence of gaseous oxygen
(GOX). All the tested formulations enhanced regression rate. Coating agents having
fluorine as a component were found to yield advantages under the test conditions [17].
The disadvantage of low regression rate of solid fuel has attracted the attention of many
researchers and means to enhance the regression rate are being tried through various routes.
Study in [9] examines the effect of the design of the fuel grain on performance
characteristics of hybrid rockets. Taking into account pressure loss due to stagnation, the
model includes a throttle component. The design is evaluated with hydroxyl-terminated
polybutadiene (HTPB) fuel using both hydrogen peroxide at 90% and liquid oxygen as
oxidizers. The aim is to develop a hybrid booster capable of the Titan 34D lift operations.
Liquid oxygen is shown to be a more effective oxidizing agent in the system than hydrogen
peroxide. Karabeyoglu et al. [10] details a two-stage low-earth orbit (LEO) rocket vehicle
launched in conjunction with the Canberra air platform. The design incorporates paraffinfueled hybrid rocket motors for first and second stages. The system is shown to handle a
31 kg payload, lifting it into a 500 km orbit. The paraffin-fueled hybrids offer safety and
reduced costs.
The studies to enhance the regression rate by changing the injection pattern of the liquid
propellant are as follows. There are two methods reported in the literature. The first method
is that of FIGG, proposed by the Propulsion Directorate of the US Air Force for its
development of hybrid propulsion for tactical missiles. And, this has been previously dealt
under section USAF.
32 Literature Review | Team ISSH
The second method is known as vortex hybrid method, reported by Orbital Technologies
[18]. In this, rather than using a wagon-wheel design having a large number of ports for
Combustion, the concept uses a co-rotating, counter-flowing combusting vortex in a
cylindrical, single central port of the fuel grain. The nozzle-end injected oxidizer spirals up
along the wall towards the head-end, causing much accelerated regression, and then
migrates to the port centre, where it spirals down towards the nozzle end and exits. For this
enhanced regression rate, one has to meet with enhanced stagnation-pressure loss in the
combustion chamber. Also, the exiting nozzle flow may have some vortices that may give
a wanted/unwanted spin to the rocket. The vortex hybrid concept is said to increase the fuel
regression rate by up to a factor of 10.
33 Literature Review | Team ISSH
Chapter 3:
Design
Methodology
3.1 Design Process Flow
The approach developed by our supervisor, Dr. Bilal is summarized in the following flow
chart, Figure 3. In Section 3.2, we will define the parameters listed in the chart.
3.2 Design Parameters
3.2.1 Burning parameters
The main burn parameter to be defined is the oxidizer-to-fuel ratio (OF).
The oxidizer fuel ratio can be given by;
OF ratio =
𝑀𝑂2 𝑁𝑂2
𝑀𝐶25 𝐻
52
𝑁𝐶25 𝐻
52
Oxidizer fuel ratio is found to be 3.45.
Theoretical specific impulse can be obtained through the dividing exit velocity and
acceleration due to gravity, which comes out to be 222.8 sec.
𝑉𝑒 = 𝑔0 𝐼𝑠𝑝
34 Design Methodology | Team ISSH
1
Input desired burn time
and average thrust
Find area of
injector Ainj
Select Chamber Pressure and
Propellants
Find diameter
of injector dinj
Calculate mass flow of
propellants 𝑚̇ 𝑝
Select port
diameter dp
Calculate Mass flow rate of fuel
ṁf and oxidizer mox
̇
Combustion
chamber pressure
Calculate Mass of fuel mf
Pc
Mass of oxidizer mox
Design
the
nozzle At, Ae,Tc
Select Oxidizer tank pressure
Mechanical
various components
Volume of oxygen tank
1
Design
Figure 3 Flow
Chart of Design
Process
35 Design Methodology | Team ISSH
End
of
Mass flow rate of propellant can be found out by dividing thrust with exit velocity
𝑚𝑝̇ =
𝑇
𝑉𝑒
From the mass flow rate of propellant and OF ratio the flow rate of oxidizer and fuel can
be calculated.
𝑚𝑓 ̇ =
𝑚𝑝̇
1+𝑂𝐹
moẋ = ṁf x OF
For the mass of fuel and oxidizer, the product of flow rate and burn time will provide us
both the masses.
𝑚𝑓 = mf ̇ x 𝑡𝑏
𝑚𝑜𝑥 = 𝑚𝑜𝑥
̇ 𝑥 𝑡𝑏
To initiate testing we need to know the pressure of oxidizer stored in the tank. Currently
we have chosen it to be 1500 psi. Now the density and volume of oxygen can be calculated.
𝐾𝐽
𝑅𝑜𝑥 = 0.2598 𝑘𝑔 °𝐾 at 300k
𝜌𝑜𝑥 =
𝑃𝑜𝑥
𝑅𝑜𝑥 𝑇𝑜𝑥
𝑉𝑜𝑥 𝑡𝑎𝑛𝑘 =
36 Design Methodology | Team ISSH
𝑚𝑜𝑥
𝜌𝑜𝑥
3.2.2 Injector and exhaust parameters
After getting burning parameters the next step is to find the combustion chamber pressure.
We have a tubular fuel grain and the burning is assumed to be uniform with changes in
grain port diameter changing with respect to time.
mox
̇
𝐴𝑖𝑛𝑗 =
𝑐𝑑 𝑖𝑛𝑗 √2𝜌𝑂𝑥 𝑃𝑜𝑥 (
𝜌𝑂𝑥 +1
2
)𝜌𝑂𝑥−1
𝛾𝑜𝑥 + 1
With Cd inj= 0.61 and 𝛾𝑜𝑥 = 1.4 the area of injector and hence the diameter can be found.
4𝐴𝑖𝑛𝑗
𝑑𝑖𝑛𝑗 = √
𝜋
The area of small individual holes in injector plate is equal to area of combined holes.
n
π 2
π
dinj = ∑ d2small
4
4
i=1
n
π
∑ d2small = d2inj
4
i=1
nd2small = d2inj
dsmall =
n=(
dinj
√n
dinj
dsmall
37 Design Methodology | Team ISSH
)2
We have chosen dsmall = 1.5mm
5
n = (1.5)2 = 12
The total number of holes in the injector plate is equal to 12.
The port diameter dp can be found out by adding twice of allowance (3mm) to the diameter
of injector plate.
dp= dinj +2A
OF ratio is given by,
OF =
̇
𝑚𝑜𝑥
𝑚̇ 𝑓
=
̇
𝑚𝑜𝑥
𝜌𝑓 .𝜋. 𝐿𝑝. 0.42 (
̇
4.𝑚𝑜𝑥
)0.47 ×10−3
𝜋.𝑑 2
𝑝
Taking density as 900 kg/m3, expression becomes
𝐿𝑝 = 0.25629(𝑂𝐹)−1 (𝑑𝑝 )−0.06 (mox
̇ )0.53
The OF ratio will decrease with respect to time, but the we are assuming that it remains
constant
throughout
the
experiment.
PROPEP
software
(http://www.arocketry.net/propep.html) was used for determining the characteristics of
different propellant formulation. The results evaluated are given below:
𝐼𝑠𝑝 = 215.44 sec
𝐶 ∗ = 5777.39 m/sec
38 Design Methodology | Team ISSH
n
π
∑ d2small = d2inj
4
i=1
nd2small = d2inj
dsmall =
dinj
√n
where
𝜌 = 4.442885 kg/m3
𝑀𝑐 = 26.70246 kg/mol
𝛾𝑐 =
𝐶𝑝
= 1.204574
𝐶𝑣
𝑇𝑐 = 3753.504 °𝐾
RA= 8.314 KJ/mol°𝐾
3.2.3 Nozzle Geometry
With the help of values from PROPEP we can easily find out the nozzle geometry. With
drag coefficient taken as 0.99, considering that throat area and diameter are at critical
conditions, for an isentropic flow:
𝐴𝑡 =
39 Design Methodology | Team ISSH
𝑚̇ 𝑝
Dt = √
4𝐴𝑡
𝜋
The chamber pressure is given by
𝑃𝑐 =
𝜌𝑐 𝑅𝐴 𝑇𝑐
𝑀𝑐
Chamber pressure would determine expansion ratio and exit area of nozzle. It is one of the
most important parameter of engine; any inaccuracy may result in failure of wall which
will result in financial loss.
𝜀=
𝐴𝑒
=
𝐴𝑡
√𝛾𝑐 (
2 𝛾𝛾𝑐𝑐+1
) −1
𝛾𝑐 + 1
2
𝛾𝑐 +1
√ 2𝛾𝑐 (𝑃𝑒 )𝛾𝑐 (1 − (𝑃𝑒 ) 𝛾𝑐 )
𝛾𝑐 − 1 𝑃𝑐
𝑃𝑐
With exit area, the exit diameter of nozzle can be calculated.
4𝐴𝑒
𝐷𝑒 = √
𝜋
The length of convergent and divergent portion of nozzle can be found by the following
expressions.
𝐿𝑐𝑜𝑛 =
1
(𝐷 − 𝐷𝑡 )
tan 𝛽 𝑒
40 Design Methodology | Team ISSH
𝐿𝑑𝑖𝑣 =
𝐷
(√𝜀 − 1) 2𝑡 + 0.7𝐷𝑡 (sec 𝜃 − 1)
tan 𝜃
With 𝛽 − 𝑐𝑜𝑛𝑣𝑒𝑟𝑔𝑒𝑛𝑡 𝑎𝑛𝑔𝑙𝑒 = 30° and 𝛼 − 𝑑𝑖𝑣𝑒𝑟𝑔𝑒𝑛𝑡 𝑎𝑛𝑔𝑙𝑒 = 15° the sum of length
of divergent and convergent portion will give the total length of nozzle.
𝐿𝑡𝑜𝑡 𝑛𝑜𝑧𝑧𝑙𝑒 = 𝐿𝑑𝑖𝑣 + 𝐿𝑐𝑜𝑛
3.2.4 Combustion Chamber Wall thickness
For wall thickness of combustion chamber we need to find the internal diameter first.
Volume of fuel can be found by the density, mass and volume relation.
𝑣𝑓 =
𝑑𝑖 = √
𝑚
𝜌𝑓
4𝑣𝑓
+ 𝑑𝑝 2
𝜋𝐿𝑝
Design pressure is the product of factor of safety and chamber pressure.
𝑃𝑑 = 𝑆𝑓 𝑃𝑐
The factor of safety taken is to be 4.
Now chamber wall thickness can be easily calculated.
41 Design Methodology | Team ISSH
𝑑𝑜 = 𝑑𝑖 + 2𝑡𝑤
3.3 HRE Design Software
Our advisor Dr. Bilal developed a design code in MATLAB ® which is his intellectual
property (IP) and being filed for patent and commercialization. We used this code in part
to validate our calculations. It will be further used for optimization. Some views of the code
are shown as follows.
Figure 4 HRE Design Program, code sample. Copyrights Dr. Bilal 2016
42 Design Methodology | Team ISSH
Figure 5 Sample Output of HRE Design Program, Copyrights Dr. Bilal 2016
3.4 Mechanical Design of Rocket Components
3.4.1 Combustion chamber
Initially we selected a hollow cylindrical combustion chamber of aluminum having
external and internal diameter of 130mm and 90mm simultaneously, and on both ends of
combustion chamber 8 holes of 8mm are made for bolts to join with other parts. As shown
in the following figure:
43 Design Methodology | Team ISSH
Figure 6 Combustion Chamber (all dimensions in mm)
Later we made changes in the design of the combustion chamber and instate of forming
bolt head on each side we form threads on each side for better fixing and changed the
internal and external diameter to 90mm and 102mm and of 392mm as shown in the
following figure :
44 Design Methodology | Team ISSH
Figure 7: Final Combustion Chamber
3.4.2 Injector Bottom Retainer
Injector bottom retainer was first designed such that to fit with the combustion chamber
perfectly by having holes for the bolts and center space to hold the injector as shown in the
following figure:
45 Design Methodology | Team ISSH
Figure 8 Injector Bottom Retainer (all dimensions in mm)
Later the design was changed as the design of combustion chamber was changed having
threaded end to fit with the new design of combustion chamber and 4 holes of 8mm around
the injector space for bolts to join it with top injector retainer as shown in the following
figure:
46 Design Methodology | Team ISSH
Figure 9: Final Injector Bottom Retainer
3.4.3 Injector Top Retainer
Injector top retainer as shown in the picture below was perfect for the initial design of the
parts it has to cylindrical pipe for the oxidizer to enter into a combustion chamber through
an injector.
47 Design Methodology | Team ISSH
Figure 10 Injector Top Retainer
Later the design was changed acording the new design of bottom retainer it has same
center hollow pipe having external diameter 17mm for oxidizer to enter and 4 holes same
as in bottom retainer for bolts and it has plate diameter of 76mm and has widht of 6mm
to sit perfectly in the space of bottom retainer as shown in the following figure .
Figure 11: Final Injector Top Retainer
48 Design Methodology | Team ISSH
3.4.4 Injector
Injector was initially designed as shown in the picture below it consists of 12 holes of
12mm diameter at an angle to create turbulence in the flow of oxidizer for better
combustion.
Figure 12 Injector (all dimensions in mm)
Later the design of an injector changed with center hole of 3mm diameter and 8 holes of
1.5mm around the center hole for axial flow of oxidizer because the machining was not
possible on such small area, but still creating enough turbulence for combustion. The
design of an injector is shown in the following figure:
49 Design Methodology | Team ISSH
Figure 13 Final Injector
3.4.5 Nozzle
Convergent-divergent nozzle was designed to produce maximum K.E for maximum thrust
with convergent angle of 15° and divergent angle of 7.5° and it has the external diameter
of 50mm for convergent part of nozzle and 90mm external diameter for divergent nozzle.
The design of a nozzle is shown in the figures below:
Figure 14 Convergent Nozzle (all dimensions in mm)
50 Design Methodology | Team ISSH
Figure 15: Nozzle
3.4.6 Nozzle Retainer
Nozzle retainer shown in the picture below was perfect fit for the initial design of and
combustion chamber having plate to fit in bolts and center space of external and internal
diameter of 50mm and 38mm to hold the nozzle in place within combustion chamber.
51 Design Methodology | Team ISSH
Figure 16 Nozzle Retainer (all dimensions in mm)
Later the design was change with threaded bottom as shown in the following figure:
Figure 17 Final Nozzle Retainer
52 Design Methodology | Team ISSH
Chapter 4:
Manufacturing of
Designed Rocket
4.1 Manufacturing and Instrumentation Costs
Projected costs of manufacturing the designed rocket and its instrumentation is dependent
on various factors, such a socio-economic milieu, availability, variable taxation etc. The
costs borne by the project are mentioned in Table 3.
Table 3 Costs of Manufacturing
Items
Cost (Pkr)
Oxidizer
15000
Solid Fuel
6000
Combustion Chamber
10000
Nozzle
15000
Valves
5000
Miscellaneous
20000
Test Bench
20000
Total
91000
Manufacturing processes are the steps through which raw materials are transformed into a
final product. The manufacturing process begins with the creation of the materials from
which the design is made. These materials are then modified through manufacturing
processes to become the required part.
53 Manufacturing of Designed Rocket | Team ISSH
4.2 Manufacturing Details
4.2.1
i.
Combustion Chamber
We started by buying a initially hollow cylinder of aluminum which was 410 mm
long, having a internal and external diameter of 80 mm and 110 mm
simultaneously.
ii.
We enclosed the cylinder into the chuck of lathe machine to start the machining of
the chamber to our desired dimensions shown in the following figure :
Figure 18 Combustion Chamber design (all dimensions in mm)
iii.
We started by setting the facing tool to the center of the cylinder and started doing
machining by doing facing on one side in two steps by removing 1 mm material
in each step.
54 Manufacturing of Designed Rocket | Team ISSH
iv.
With the same tool, we did turning process on the cylinder to achieve the desired
outer diameter of 102 mm; we did this process in 4 steps to 393 mm along the
length of cylinder by removing 2mm material in each step.
v.
After the turning process, we changed the tool to one edge cutting tool for boring
process to get the required internal diameter of 90mm, we did this process in 5
steps moving tool 393mm along the length of cylinder and 2mm width wise to
remove 2mm material in first 4 steps and 1.5mm in last step leaving the space for
finishing; after that rotating the work piece with same speed and rubbing the fine
sand paper inside the hole to get rid of rough surface.
vi.
To form 1 mm threads inside cylinder to about 25.4 mm long we changed the tool
to 60° degree cut tool and change the work settings of lathe machine for 1mm
threads according to the manual.
vii.
After the forming the threads the we switch the side of our work piece to work on
other side and complete making our chamber; we changed settings back to normal
and the tool to facing tool to remove extra part that was enclosed in a chuck we
removed the extra part that was 15 mm long with the help of facing tool.
viii.
After that we did facing and removed 1mm material to get the desired length and
remove the roughness.
ix.
We repeated step 6 to get threads on the other side of cylinder.
x.
Final combustion chamber was 392mm long and has internal and external
diameter of 90mm and 102mm simultaneously; which was our design
requirement.
55 Manufacturing of Designed Rocket | Team ISSH
Figure 19 Manufactured Combustion Chamber
4.2.2 Nozzle
Nozzle is used to increase the velocity of the exhaust gasses at high temperature and low
pressure to provide maximum thrust.
We are using convergent-divergent nozzle.
4.2.2.1 Manufacturing
i.
We purchased graphite cylindrical part; which was initially 200mm long and has
diameter of 100mm, we have to achieve the design and dimensions of the part as
shown in figure below:
56 Manufacturing of Designed Rocket | Team ISSH
Figure 20 Nozzle design (all dimensions in mm)
ii.
We enclosed the part into the chuck of lathe machine and after setting them to the
center of the part with tool; we started doing facing, it was done in 2 steps; in each
step 1mm material was removed.
iii.
After facing using the same tool we started with turning process along the length
of 177mm in 5 steps removing 2mm material in each step.
iv.
For next step, we shifted the side and removed the extra length of 21mm and did
facing of 1mm for finishing.
v.
We shifted the part to drilling machine to form through hole of 14mm.
vi.
Again, shifting the part to lathe we started doing turning along the of 70mm in 2
steps removing 2mm material in each step; then moving along the length of
57 Manufacturing of Designed Rocket | Team ISSH
44mm in 8 steps removing 5mm material in first 5 steps and 2mm material in last
3 steps.
vii.
For the internal profile of divergent part of nozzle changed the tool to 7.5° angle
tapered reamer and started reaming process along the of 15mm and achieved the
desired shape.
viii.
And for the internal profile of convergent part of nozzle shifted the side of part
and repeat the reaming process along the length of 132mm by using 15° angle
tapered reamer.
ix.
Rotating part at low speed and rubbing fine sand paper inside the part to remove
the roughness.
Figure 21 Manufactured Nozzle
4.2.3 Injector
Injector is used to introduce the oxidizer into the combustion chamber. Design and
dimension of an injector is shown in the following figure.
58 Manufacturing of Designed Rocket | Team ISSH
Figure 22 Injector design (all dimensions in mm)
i.
Started with cylindrical piece with initial length of 24mm and diameter of 40mm.
ii.
We enclosed the part into the chuck of lathe machine. After aligning the workpiece with tool, we started doing facing, it was done in 2 steps; in each step 1mm
material was removed.
iii.
After facing using the same tool we started with turning process along the length
of 13mm in 3 steps removing 2mm material in first 2 step and 1mm in last step.
Then continuing turning process along the length of 6mm reducing diameter to
59 Manufacturing of Designed Rocket | Team ISSH
17mm in 5 steps removing 4mm material in first 4 steps and 2mm material in last
step.
iv.
Using drill bit of 3mm drilled the through hole of 3mm.
v.
After the drill again shifting side of the work piece removing extra length of 9
mm with parting tool.
vi.
After removing extra length, we did facing and removed 1mm material.
vii.
Using same facing tool, we the parting of 6mm and 2mm deep leaving 17mm
from center in 3 steps removing 2mm in each step.
viii.
Using boring tool, we increased the diameter to 16mm along the length of 6mm in
7 steps removing 2mm material in 6 step and 1mm is last step.
ix.
After boring changing the side of work piece and using the drill bit of 1.5mm 8
through holes are drilled at an angle of 45° from each other and 3mm away from
center of center hole.
Figure 23 Manufactured Injector
60 Manufacturing of Designed Rocket | Team ISSH
4.2.4 Retainers
Retainers are used to hold the parts on its place.
4.2.4.1 Injector Top Retainer
i.
We started with the cylinder 80mm in diameter and 50mm long; the required
design and dimensions of a retainer is shown in the picture below:
Figure 24 Injector Top Retainer design (all dimensions in mm, degrees)
61 Manufacturing of Designed Rocket | Team ISSH
ii.
We enclosed the part into the chuck of lathe machine and after setting the center
of the part with tool; we started doing facing, it was done in 2 steps; in each step
1mm material was removed.
iii.
Using same tool turning is done on the work piece along the length of 40mm and
reducing the diameter of part to 13mm in 15 steps removing 10mm in first 5 steps
and removing 4mm in further 5 steps and removing 2mm in last 5 steps for
smooth finishing.
iv.
After turning using the drill bit of 5mm drilled the through hole.
v.
Changing the side and using the facing tool; repeated the facing process in 2 steps
removing 1mm material in each step.
vi.
Using same tool repeated the turning process along the length of 6mm in 2 steps
removing 2mm material in each step.
vii.
After turning parting of 6mm is made; away from 17mm to the center of work
piece and 2mm deep in 3 steps removing 2mm in each step.
viii.
Using the drill bit of 8mm; 4 through holes are drilled at an angle 90° from each
other and 25mm away from center of center hole.
62 Manufacturing of Designed Rocket | Team ISSH
Figure 25 Manufactured Injector Top Retainer
4.2.4.2 Injector Bottom Retainer
i.
Started by using the cylindrical work piece of 110mm diameter and 50mm long;
the required design and dimension are shown in Figure 26.
ii.
We enclosed the part into the chuck of lathe machine and after setting the center
of the part with tool; we started doing facing, it was done in 2 steps; in each step
2mm material was removed.
iii.
Using same tool turning is done on the work piece along the length of 40mm and
reducing the diameter of part to 103mm in 4 steps removing 2mm material in each
step.
iv.
After turning using the drill bit of 8mm, 5 through holes are drilled; 1 at the center
of work piece and remaining 4 holes are drilled 25mm away from the center of
center hole and at an angle of 90° from each other.
63 Manufacturing of Designed Rocket | Team ISSH
Figure 26 Nozzle Retainer design (all dimensions in mm)
v.
Again, setting the center of tool with the work piece and using boring tool
increases the diameter center hole to 16mm in 4 steps removing 2mm material in
each step.
vi.
Continuing with boring process along the length of 26mm; increased the diameter
to 35mm in 5 steps removing 5mm in first 3 steps and removing 2mm material in
last 2 steps. Then along the length of 20mm; increased the diameter to 76mm in 7
steps removing 10mm material in first 3 steps then removing 5mm material in 1
step and then removing 2mm material in last 3 steps.
64 Manufacturing of Designed Rocket | Team ISSH
vii.
Switching the side of work piece and using facing; done facing in 4 steps
removing 5mm material in first 2 steps and 2mm material in last 2 steps.
viii.
Using threading tool and setting automatic settings for 1mm threads along the
length of 27.23mm threading is done.
Figure 27 Manufactured Injector Bottom Retainer
Figure 28 Assembled top and bottom of retainer with bolts
4.2.4.3 Nozzle Retainer
i.
Started by using the cylindrical work piece of 110mm diameter and 80mm long;
the required design and dimension are shown in the following figure :
65 Manufacturing of Designed Rocket | Team ISSH
Figure 29 Nozzle Retainer design (all dimensions in mm)
ii.
After enclosing the part into the chuck of lathe machine and setting the center of
the part with tool; started doing facing, it was done in 2 steps removing 1mm
material in each step.
iii.
Using same tool turning is done on the work piece along the length of 71mm and
reducing the diameter of part to 102mm in 4 steps removing 2mm material in each
step; and further reduced the diameter to 58mm along the length of 38mm in 12
steps removing 5mm material in first 8 steps and 2mm material in last 4 steps.
iv.
After turning using the drill bit of 8mm, drilled the through hole.
v.
Then using boring tool increased the diameter center hole to 50mm in 9 steps
removing 5mm material in first 8 steps and 2mm in last step.
66 Manufacturing of Designed Rocket | Team ISSH
vi.
Changing the side and continuing with boring process along the length of 27mm;
increased the diameter to 86mm in 9 steps removing 5mm in first 6 steps and
removing 2mm material in last 3 steps.
vii.
Using the facing tool and repeating the facing process in 4 steps removing 2mm
material in first 3 steps and 1mm material in last step.
viii.
Using threading tool and setting automatic settings for 1mm threads along the
length of 27mm threading is done.
Figure 30 Manufactured Nozzle Retainer
4.3 Fuel Grain Casting
There were different choices of waxes available in the market. The wax chosen by
Team ISSH was Pentacosane having chemical formula of C25 H52, a product of BASF
Chemical Company with following properties
i.
Melting Point : 54°C
67 Manufacturing of Designed Rocket | Team ISSH
ii.
Boiling Point : 401°C
iii.
Density : 0.801 g/ml
iv.
Molar Mass : 352.69 g/mol
The chemical name and properties were a result of series of tests carried out at HEC
Karachi University.
Figure 31 Pentacosane packed in bags with rubber sheet and other apparatus used for casting
The team has two choices for melting of wax, heated bath or through direct heating.
Heated bath was a time consuming procedure so direct heating was chosen 1.5 kg of wax
was casted into the rubber sheet sealed through adhesive to prevent leakage. Required
rubber sheet was fitted into the chamber, circular disks were used which helped in the
process.
5% of carbon was added into the wax during heating to increase the amount of impurities
of wax which would help increase the melting and boiling point.
68 Manufacturing of Designed Rocket | Team ISSH
Figure 32 Casting process in progress
Figure 33 Molten wax being poured into the combustion chamber through funnel
The fuel grain was casted by wax imported from Germany, manufactured by BASF which
has melting point of about 400°C. The wax was casted in a rubber sheet which was meant
to provide insulation to reduce the direct heat transfer to the combustion chamber wall,
resulting in safety.
69 Manufacturing of Designed Rocket | Team ISSH
Figure 34 Combustion chamber view after casting
Figure 35 Rubber sheet acting as insulator
70 Manufacturing of Designed Rocket | Team ISSH
Figure 36 Fuel grain fitted inside the combustion chamber
The team tried to remove the shrinkage effect from the grain by re-casting wax into the
rubber section but the defect was not eliminated completely.
71 Manufacturing of Designed Rocket | Team ISSH
Chapter 5:
User Manual with
Designed HRE Shaheen
This kit contains one complete set of components for use in a Shaheen rocket motor
only, no other hardware can be used with this reload.
Figure 37 Shaheen Hybrid Rocket Engine Schematic
5.1 Engine Specifications
Table 4 Engine Specification
Total
Peak
Burn
Oxidizer tank
impulse
thrust
time
pressure
(s)
(N)
(s)
(psi)
222
1000
10
1500
5.1.1 Caution/Warning
i.
Do not open reload until ready to use.
72 User Manual with Designed HRE Shaheen | Team ISSH
ii.
Correct parameters must be used according to the table above; otherwise motor
failure may occur at ignition or during flight.
iii.
Please be sure to read and understand all instructions in this manual prior to
assembly. Make sure to receive all the parts required for reload.
iv.
Assemble the Rocket Motor according to the directions only. Do not modify any
part of the reload.
v.
Combustion Chamber and Nozzle may be hot after firing. Please use extreme
caution when handling or disassembling recently fired motors.
5.2 Package Part List
5.2.1 Motor Hardware
i.
Oxidizer Tank Assembly
ii.
Combustion Chamber
iii.
Nozzle
iv.
Fuel Grain
5.2.2 Hybrid Reload
i.
Hybrid Fuel Grain
ii.
Instruction Manual
iii.
Safety Igniter
73 User Manual with Designed HRE Shaheen | Team ISSH
Figure 38 Oxidizer Tank
Figure 39 Fuel Grain
74 User Manual with Designed HRE Shaheen | Team ISSH
5.3 Assembling Procedure
Step 1: Begin with a clean, oil free motor case. Start by installing the push lock injectors
into the bottom of the oxidizer tank assembly.
Step 2: Mount the LOX assembly on the oxygen tank. Attach the oxygen line to the
regulator. The fitting tightening instructions must be followed well. It is recommended
that the oxygen bottle be laid on its side to prevent damage to the valve or regulator in
case it is knocked over. The aluminum hose fittings are also particularly susceptible to
this type of damage. Remember that oxygen is stored as a compressed gas and thus tank
orientation is unimportant.
Figure 40 Condi Nozzle
75 User Manual with Designed HRE Shaheen | Team ISSH
Figure 41 Nozzle Divergent portion
Figure 42 Combustion Chamber
Step 3: Fully open the oxygen tank valve and adjust the regulator to between 80 and 100
psi. If you hear any leakage, turn the oxygen off and make sure the fittings are tightened.
76 User Manual with Designed HRE Shaheen | Team ISSH
Step 4: Slide the combustion chamber section over the injector section
of the LOX tank, and insert the retaining bolts. The bolts should be
inserted to just below the depth of the tube. Do not over tighten the
bolts.
Step 5: Slide the fuel grain into the combustion chamber.
Step 6: Slide the Nozzle into the Combustion Chamber and align the
retaining bolt holes carefully.
Step 7: Attach the battery cables to a 12V DC source capable of supplying at least 10
amps. Either a car battery or a gel-cell battery will be adequate.
Step 8: Double check that the ignition wires are clear of any part of the motor or other
materials, and especially away from you! Do not place any part of your body near the
ignition circuitry while this test is in progress – serious electric shock could result.
Step 9: Make sure that the launch pad area is clear prior to checking the oxidizer fill
system.
Step 10: once you have ensured that the system is setup your rocket is ready, LOX supply
is connected and you are ready to fill your motor connect the igniter leads to the 24 volts
launch system.
For throttling the engine, Team ISSH used FrSky Taranis X9D Plus Transmitter which
remotely operated an servo-actuated ball valve developed by AeroX.
77 User Manual with Designed HRE Shaheen | Team ISSH
Figure 43 FrSky Taranis X9D Plus Transmitter
5.4 Post Firing Cleaning Procedures
Step 1: Once the process is completed you can begin with the
disassembly and cleaning process.
Step 2: We recommend you removing all retention bolts in the
combustion chamber only! Do not disassemble the oxygen tank for
any reason.
Step 3: Pull the Combustion chamber away from the Oxidizer tank, and
Nozzle. Push any remaining fuel and liner from the Combustion
Chamber and Clean with Soap and Water.
Note: Other Solvents Such as Lighter Fluid, Acetone and Vinegar have
been known to aid in the cleaning process
78 User Manual with Designed HRE Shaheen | Team ISSH
Step 4: Clean any remaining oils or lubricants from the outside of the
nozzle and Injector Baffle. When cleaning the nozzle, leave any built
up char, ash, or carbon deposits on the inside of the Nozzle Bell. This
is a normal buildup and will keep your nozzle lasting a long time, by
providing an ablative layer during your next burn.
Step 5: Dry all Parts of the motor, and store for next use
Step 6: When done with your motor for the day, wipe all parts down with WD-40. This
will keep your parts looking like new, resistant from fading, fingerprinting and corrosion.
Be Sure NOT to get and WD-40 in the Oxidizer tank!!
5.5 Safety and First Aid
Hybrid rocket Motor Reloads will not burn without the presence of a high temp heat source,
and strong oxidizer. If for some reason, any part of a reload is ingested, induce vomiting
and seek medical attention.
5.6 Disclaimer
Our rocket engine ensures reasonable care during the design and manufacture process.
Because we cannot control the use or storage of our products, our engine cannot be held
responsible for any personal injury or property damage resulting from the handling, use or
storage of its products.
79 User Manual with Designed HRE Shaheen | Team ISSH
Chapter 6:
Experimental
Results and Discussion
Shaheen HRE was test fired twice to test design parameters and motor performance. The
tests were fully automated and appropriately instrumented. The test matrix is given below.
Table 5 Test Fire Matrix
Test
Testing Mode
Test Stand
Date
Remarks
Static
Test stand bolted in 15-11-2016
Successful
place
premature
No.
1.
fire;
shutdown
2.
Static
Test on purpose build
Successful
test;
vehicle; anchored
longest duration
6.1 Test Procedure
1. Ignition System. Rocket will be ignited through electrical igniters.
2. Misfires. If engine fails to ignite through electric ignition, then the battery will be
disconnected and oxygen tank valve will be closed.
3. Launch Safety. Countdown will be used, and it would be ensured that everyone
is attentive and is at safe distance.
80 Experimental Results and Discussion | Team ISSH
4. Testing site. Testing will be carried out in open atmosphere. Testing bench will
be secured to the ground. It will be ensured that no dry grass is nearby testing site.
5. Test fire safety. Any flammable or explosive material would be placed near the
engine
6. Weather conditions. Testing will done with wind speeds no greater than 10 miles
per hour. It will be ensured that no dry grass is nearby testing site.
6.1.1 Assembling the engine
Fuel grain is placed cautiously in the chamber and no press-fitting is done when tightening
caps which may damage grooves. Damaging of grooves may result in leakage of oxidizer
from the chamber.
6.1.2 Connecting the Igniter
The testing area is evacuated, with the team members present near the engine. Fire
extinguisher and first aid equipment will be present in case of emergency. Team will ensure
to perform two tasks: connecting filling lines to tanks (which will be closed) and
connecting the igniter to engine. The opening of oxidizer valve will be the last step.
6.1.3 T-0:00
Motor is ready for firing. Team awaits the final countdown; all the members and
audience will be present at a safe distance. Safety officer will confirm that range is safe.
81 Experimental Results and Discussion | Team ISSH
Countdown begins for fire.
“Initiating ignition in five, four, three…”
6.1.4 T-5 seconds
Five seconds before the launch, the button is pressed. In case of unforeseen circumstances
safety officer will just need to disconnect the fuel line and combustion chamber.
6.1.5 T-4 seconds
Valve is opened and internal ignition starts, loss of signal to igniter would abort the test.
This is done in case of undesirable ignition.
6.1.6 T-3 seconds
Igniter is fired. Small pyrotechnic charge ignites the initial flow of oxygen which helps to
preheat the combustion chamber to preheat before main valve is opened and oxygen is
drown in the combustion chamber. Preheating the chamber will help proper combustion of
oxidizer, decreasing the chances of suffocation inside fuel grain.
6.1.7 T +0 seconds Launch
The main valve opens and motor fires. Required thrust of achieved within burn time. The
engine keeps in contact on the test stand and the data is sent to the ground station. The abort
system is active for 10 seconds after burnout. Oxidizer valve is closed, fuel line is
disconnected and engine is left in open air to cool down.
82 Experimental Results and Discussion | Team ISSH
6.2 First Static Test
Team ISSH conducted its first static test on 15th November, 2016 under the supervision of
Dr. Bilal, our advisor. The oxidizer tank used for the test was of 12 liters having pressure
oxygen gas at a pressure of 1000 psi.
Figure 44 Oxidizer tank with needle valve regulator
83 Experimental Results and Discussion | Team ISSH
Ignition was smooth but as the pressure started to increase the mounts holding the engine
to test bench lost grip and oxidizer supply got disconnected which was not clamped
properly. The total burn time of the test was 2.5 sec.
Figure 45 First static test in progress
6.2.1 Post Test Visual Inspection after First Test
The rocket components were visually inspected after the test to analyze burn performance.
6.2.1.1 Injector Plate
Injector plate was not affected after the testing because heat was not transferred to the
oxidizer tank, if there was transfer of heat or increase of back pressure this would result in
a failure and explosion. The carbon was easily removed by a piece of cloth. The caps were
also cleaned from the carbon after the test.
84 Experimental Results and Discussion | Team ISSH
Figure 46 Injector Plate (back view, post test)
Figure 47 Injector Plate (front view, post test)
Figure 48 Injector Top Retainer (post test)
85 Experimental Results and Discussion | Team ISSH
6.2.1.2 Fuel Grain
The fuel grain was removed from the rocket and dissected for understanding the burn
pattern. The dissected view, shown in Figure 49 Fuel Grain Single Section after 1st testFigure 49Figure 50. It clearly shows asymmetric burning and burn through from pre-existing cracks
in the grain. X-Ray radiographical images (Figure 51) of the grain further shows the
substructure of the tested fuel grain, and existence of air voids in the fuel grain. This helped
us avoid this mistake in carefully casting the fuel grain for the next test. Remarkably, there
was no visible damage to any other component. Even the insulation was intact after the
test.
Figure 49 Fuel Grain Single Section after 1st test
86 Experimental Results and Discussion | Team ISSH
Figure 50 Complete Fuel Grain section view after 1st test
6.2.1.3 Test Stand
The thrust was so excessive that the 100 kg load cell was saturated and we could not get
accurate thrust readings. However it was definitely more than 1000 N, which can also be
seen that the load cell was bent permanently out of shape and the test bench had to be
abandoned for repair. This may have been caused by voids in the fuel grain and asymmetric
burning. Thankfully, the premature shutdown prevented a potential motor burn-through.
This may also be the reason.
6.3 Second Static Test
The second test was conducted on 17th December, 2016; mounting was done on the chassis
of car, designed and manufactured by another group of students working under Dr. Bilal
87 Experimental Results and Discussion | Team ISSH
(namely Khizar, Waqas, Mohsin and Umer). The chassis was placed in an underground pit
due to safety and calibration. The thrust could not be calculated properly due to improper
calibration. The burn time for the fuel was improved to 5 seconds.
Voids
Figure 51 X-Ray of Fuel Grain (post test)
The engine was designed to achieve burn time of more than 10 sec, testing engineer present
at the site tried to throttle the engine at a high rate due to which the oxidizer valve closed
accidently. This break up in the oxidizer resulted in reduced burn time.
88 Experimental Results and Discussion | Team ISSH
Load Cell
Figure 52 Visual Inspection of Test Stand (post test)
Figure 53 Test being carried out in the test pit
(With permission of Khizar et al)
89 Experimental Results and Discussion | Team ISSH
Figure 54 CAD model of car chassis mounted with oxidizer tank and engine
(with permission from Khizar et al)
Figure 55 1000 psi pressured gas can be clearly seen on the gauge
6.4 Third Static Test
This test was also conducted on 17th December, 2016. The ignition and combustion was
very smooth and the flame was very much impressive. The burn time for fuel was more
than 40 sec which was a great achievement for Team ISSH. Weighing machine fitted with
90 Experimental Results and Discussion | Team ISSH
load cell from Salter Precision Machining Inc was placed in front of chassis; the resulting
push of the car would give the applied force on the load cell. The thrust value was around
800N, with a burn time of 35 seconds.
Figure 56 3rd test being carried out in the test pit
The remaining unburnt wax of very low quantity was photographed by the team. Oxidizer
tank of 9 liters was utilized.
Figure 57 Weighing machine fitted with Salter Precision Machining Inc load cells
Specifications of the digital force measurement equipment are as mentioned below.
91 Experimental Results and Discussion | Team ISSH
Maximum weight x resolution = 180kg x 0.1kg, 28st 81lb x 1/4lb, 400lb x 0.2lb
Dimensions = 30.4 x 30.4 x 3.1cm
Batteries = 1 x CR2032
6.4.1 Post Test Visual Inspection after First Test
The rocket components were visually inspected after the test to analyze burn performance.
It was found that only sliver remained of the remaining fuel, and almost all of it was
consumed during the test. Moreover the very smooth burning made us conclude that the
fuel grain quality was of much more consistent quality.
Figure 58 Nozzle and its Retainer (after third test)
92 Experimental Results and Discussion | Team ISSH
Figure 59 Graphite Nozzle Insert with Residue Deposits
Figure 60 Inside View of Combustion Chamber with Wax and Rubber Residue
93 Experimental Results and Discussion | Team ISSH
Figure 61 Nozzle Retainer after Third Test
Figure 62 Residue (unburnt wax) after 3rd test
94 Experimental Results and Discussion | Team ISSH
Graphite nozzle was able to withstand elevated temperature of more than 1200 °K. Even
after 3 consecutive tests the equipment is still able for further testing except for the fuel
grain which needs to be casted again.
6.5 Testing Videos
The videos for our tests are uploaded on YouTube.
https://www.youtube.com/playlist?list=PLNX6y2X4XodqH_E3hoFMkXSCAhpSzAiT9
It can be seen in the videos that the rocket gives a burn time of about 40 seconds,
unprecedented for hybrid rockets in Asia, as well as a steady burn. More testing will be
carried out in the future for improving the design.
95 Experimental Results and Discussion | Team ISSH
Chapter 7:
Design
Recommendations
A 1000 N thrust hybrid rocket engine’s prototype was designed, manufactured and tested.
Many fuels and oxidizers were considered during the beginning of the project. PROPEP
software was used to calculate the exhaust product properties and the result was evaluated
and for calculations we used different books and find OF ratio, thrust coefficient, thrust,
and specific impulse. From this, throat diameter of 14 millimeters, O/F ratio of 3.14 was
chosen as the optimal design. An aluminum combustion chamber was turned to size and
threaded to accept the bottom injector retainer and nozzle retainer both of aluminum, the
nozzle was turned out of graphite to the appropriate shape, and the fuel grain was cast in a
rubber tube. Finally, the motor was instrumented to measure the chamber pressure,
generated thrust, and oxidizer mass flow rate as functions of time throughout the burn. The
design, construction, and testing of a 1000 N thrust hybrid rocket engine lies well within
the realm of an upper division undergraduate course. The entire process contained within
this project is an invaluable method by which the course content becomes fully understood
and prerequisite knowledge reinforced.
96 Design Recommendations | Team ISSH
Chapter 8:
Conclusion and
Future Work
The main objective of our project was to design and fabricate a hybrid rocket engine which
would generate a thrust of 1000N with a burn time of 10 seconds. We have completed our
manufacturing and testing, and proven a thrust level of 800N (measured) and 1500 N
(approximated), with burn time of 40 seconds. We have therefore surpassed the
expectations mentioned in the Statement of Requirements.
For future research, it is important to have more extensive instrumentation and high speed
or infrared videography for better understanding of the performance characteristics.
97 Conclusion and Future Work | Team ISSH
References
[1] G. Sutton and R. Biblarz “Rocket Propulsion Elements”, 7th Edition, John Wiley and
Sons, NY, 2002.
[2] Cengel and Boles “Thermodynamics: An Engineering Approach”,8th Edition, McGrawHill Education, NY, 2006.
[3] De Luca, L.T. et al (Ed), “Chemical Rocket Propulsion”, Springer, ISBN 978-3-3192016, 27748-6
[4] Turner, M.J.L.,“Rocket and Spacecraft Propulsion”, 2nd Edition, Springer-Verlag,
Berlin, 2009.
[5] Kohei Ozawa et al, "Static Burning Tests on a Bread Board Model of Altering-intensity
Swirling-Oxidizer-Flow-Type Hybrid Rocket Engine", 52nd AIAA/SAE/ASEE Joint
Propulsion Conference, Propulsion and Energy Forum, (AIAA 2016-4964)
[6] Barato, F., et al, “Integrated approach for hybrid rocket technology development”, Acta
Astronautica, Vol 28, pp 257-261, Nov-Dec 2016.
[7] E. Alkuam and W. Alobaidi, “Experimental and Theoretical Research Review of
Hybrid Rocket Motor Techniques and Applications”, Advances in Aerospace Science and
Technology, No 1, 2016, pp 71-82.
[8] Hudson, M.K., Shanks, R.B., Snider, D.H., Lindquist, D.M., Luchini, C. and Rooke, S.
“UV, Visible, and Infrared Spectral Emissions in Hybrid Rocket Plumes.” International
Journal of Turbo and Jet Engines, 1998, No. 15, pp. 71-87.
98 References | Team ISSH
[9] Vonderwell, D.J., Murray, I.F. and Heister, S.D. (1995) Optimization of HybridRocket-Booster Fuel-Grain Design. Journal of Spacecraft and Rockets , 32, 964-969.
[10] Karabeyoglu, A., Falconer, T., Cantwell, B. and Stevens, J. (2005) Design of an
Orbital Hybrid Rocket Vehicle Launched from Canberra Air Platform. 41st
AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, Joint Propulsion
Conferences , (American Institute of Aeronautics and Astronautics ), 10-13 July 2005, 122.
[11] Donahue, B.B. (2004) Beating the Rocket Equation: Air Launch with Advanced
Chemical Propulsion. Journal of Spacecraft and Rockets , 41, 302-309.
[12] Hudson, M.K., Shanks, R.B., Snider, D.H., Lindquist, D.M., Luchini, C. and Rooke,
S. (1998) UV, Visible, and Infrared Spectral Emissions in Hybrid Rocket Plumes.
International Journal of Turbo and Jet Engines , 15, 71-87.
[13] Wright, A.M., Foley, P., Tilahun, D., Reason, M., Bryant, C., Patton, J. and Hudson,
M.K. (2000) The Effect of High Concentration Guanidinium Azo-Tetrazolate on Thrust
and Specific Impulse of a Hybrid Rocket. 36th AIAA/ASME/SAE/ASEE Joint Propulsion
Conference and Exhibit, Joint Propulsion Conferences , (American Institute of Aeronautics
and Astronautics ), 16-19 July 2000, 1-9. http://dx.doi.org/10.2514/6.2000-3885
[14] Shanks, R. and Hudson, M.K. (2000) A Labscale Hybrid Rocket Motor for
Instrumentation Studies. Journal of Pyrotechnics , No. 11, 1-10.
99 References | Team ISSH
[15] Risha, G.A., Boyer, E., Evans, B., Kuo, K.K. and Malek, R. (2004) Characterization
of Nano-Sized Particles for Propulsion Applications. Materials Research Society
Symposium Proceedings , (Materials Research Society ), 800, 243-254.
[16] Larson, D.B., Boyer, E., Wachs, T., Kuo, K.K., DeSain, J.D., Curtiss, T.J. and Brady,
B.B. (2011) Characterization of the Performance of Paraffin/LiAlH4 Solid Fuels in a
Hybrid Rocket System. 47th AIAA/ASME/SAE/ASEE Joint Propulsion Conference &
Exhibit , Joint Propulsion Conferences , (American Institute of Aeronautics and
Astronautics ), July, 1-15.
[17] Sossi, A., Duranti, E., Manzoni, M., Paravan, C., DeLuca, L.T., Vorozhtsov, A.B.,
Lerner, M.I., Rodkevich, N.G., Gromov, A.A. and Savin, N. (2013) Combustion of HTPBBased Solid Fuels Loaded with Coated Nanoaluminum. Combustion Science and
Technology, 185, 17-36.
[18] Abel, T., "Hybrid Rockets," Aerospace America, Vol. 37, Dec. 1999, p. 75.
100 References | Team ISSH