sriram project
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DESIGN, DEVELOPMENT & TESTING OF LIQUID
PROPELLANT ROCKET ENGINE
A
PROJECT REPORT
Submitted by
ARUN KUMAR .S (70805101004)
DINESH KUMAR .B (70805101013)
PRAVEEN KOP .P (70805101707)
SRI RAMAN .A (70805101708)
in partial fulfillment for the award of the degree
of
BACHELOR OF ENGINEERING
IN
AERONAUTICAL ENGINEERING
HINDUSTHAN COLLEGE OF ENGINEERING AND
TECHNOLOGY
COIMBATORE 641 032
ANNA UNIVERSITY: CHENNAI 600 025APRIL 2009
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ANNA UNIVERSITY: CHENNAI 600 025
BONAFIDE CERTIFICATE
Certified that this project report DESIGN, DEVELOPMENT & TESTING OF
LIQUID PROPELLANT ROCKET ENGINE is the bonafide work of S.
ARUN KUMAR, B. DINESH KUMAR, P. PRAVEEN KOP, A. SRI
RAMAN who carried out the project work under my supervision.
SIGNATURE SIGNATURE
Wg Cdr Prof. Arun Kumar Adak Mr. C. Parthasarathy, M.E
PROFESSOR & LECTURER
HEAD OF THE DEPARTMENT SUPERVISOR
Department of Aeronautical Engineering Department of Aeronautical Engineering
Hindusthan College of Engg. & Tech. Hindusthan College of Engg. & Tech.
Coimbatore 641 032 Coimbatore 641 032
Submitted for University Project viva-voce examination conducted on 23 April
2009.
Internal Examiner External Examiner
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DECLARATION
I hereby declare that the entire work embodied in this dissertation has been
carried out by us and no part of it has been submitted for any degree or diploma of
any institution previously.
Signature Signature Signature Signature
S. ARUN KUMAR B. DINESH KUMAR P.PRAVEEN KOP A. SRI RAMAN
Place: Coimbatore
Date: 23 April 2009
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ACKNOWLEDGEMENT
We took up an ambitious project ofDESIGN, DEVELOPMENT & TESTING OF LIQUID
PROPELLANT ROCKET ENGINE. Initially we were wary about such a different and difficult
undertaking. However, we came across COINDIA who were in a position to help us out by manufacturing
the rocket engine of our design. When we approached them and discussed our concept with Mr. S.
Devananthan, Tool Room manager, COINDIA we were convinced that we have contacted the right
personnel who are in a position to help us out.
During the process of actual manufacture, which took up nearly two months the interactions
between Mr. S. Devananthan and ourselves dispelled all our initial doubts, because of the efforts put in by
him. Mr. S. Devananthan helped us to the maximum in every possible way and we are deeply moved byhis kindness, despite the odds that confronted us.
Our sincere and heartfelt thanks are due to him in helping us to give shape to our design and to
achieve our target.
We also express our profound gratitude to Prof. Arun Kumar Adak, Head of the Department,
Mr. D. Viswanathan, Associate Professor and our guide Mr. C. Parthasarathy and all the teaching and
non-teaching staffs of the Department of Aeronautical Engineering of our college for their suggestions
and encouragement.
Our thanks are also due to
Mr. K.T. Sitaraman, Polymer Technologist,143, Poompuhar Nagar, Thadagam Road, Edayarpalayam, Coimbatore - 641025
Mr. T. V. Sankaran, Praghashree Systems (P) Limited43/44, K.G.Layout, Sabapathy Street, K.K.Pudur, Coimbatore - 641038
Mr. Ritesh shah, Shaileshco Industries38, SIDCO Indutrial Estate, Kurichi, Coimbatore - 641021
Mr. Venkatesh, COINDIA CAD solutions Mr. Kumar, Chief CNC Engineer, COINDIA Mr. Anand, CNC program setter, COINDIA Mr. Vimal raj, Design Engineer, COINDIA
340-342, Avarampalayam Road, K.R. Puram, Coimbatore - 641006
Mr. M.Kumaresan, Best Forgings India (P) Limited15-A, private Industrial Estate, Coimbatore - 641021
Mr. Jayaraman, Weightronic systems479B,Avinashi Road, Near Suguna Kalyana Mandapam, Peelamedu, Coimbatore-641004
Mr. K. Santhosh Kumar, EdgeCAM specialist23/3 (Old No. 10/3) Archana apartments, Sarangapani Street, T.Nagar, Chennai600017
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Heatran Engineering CorporationNo.199, Dr. Nanjappa Road, Coimbatore-641018
Moiz & CompanyDr. Nanjappa Road, Coimbatore 641018
For helping us out in achieving the target.
Finally we express our sincere thanks to all who have helped us throughout the project.
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Abstract
Satellites play a vital role in the development of a country. Depending upon the dimensions and
weight of a satellite the launch vehicle and hence the engine has to be properly designed. Engines have to
be designed in such a way that they should give maximum power so as to lift off from the surface. But
lifting off from the surface doesnt make it all; it has to reach the cosmos and must be able to inject the
satellite into the orbit with exact orbital velocity. Conventional launchers and engines are complicated
involving huge expenses and efforts, and can inject huge satellites into orbit. But incase if a small satellite
is designed and has to be tested, it is not practical to launch it immediately into the orbit with the existing
huge satellite launch vehicles because of the high cost. Also in case of failures huge amount of money and
precious time will be lost. In this situation a small satellite launch vehicle (SSLV) with a compatible
engine will nearly rectify all these drawbacks. So we as a group wanted to develop a technology suitable
for launching a payload of about 50 Kg into a Low Earth Orbit (LEO). Our work here is to design,
manufacture and to test a 500 Kgf thrust producing liquid propellant rocket engine. This will be ourindigenous model since all the works have been done using the materials which are locally available. We
believe depending upon the success it will help our nation to develop in this field.
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TABLE OF CONTENTS
ACKNOWLEDGEMENT i
ABSTRACT ii
LIST OF TABLES vi
LIST OF FIGURES vi
Chapter I INTRODUCTION 2
1.1 Introduction to liquid rocket engines 2
1.2 Fundamentals of Rocket Propulsion 4
1.3 Some standards for amateur rocket designers 5
1.4 Design Section Calculation of Mass flow rates 6
Chapter II COMPRESSIBLE FLOW
2.1 Continuity Equation 9
2.2 Bernoullis Equation from Eulers equation of motion 9
2.3 Isentropic Relations 11
2.4 The Nozzle 16
2.5 Design section Design of Nozzle 18
Chapter III THE COMBUSTION CHAMBER
3.1 Combustion chamber 25
3.2 Design section Design of Combustion chamber 26
3.3 The Injector and its types 29
3.4 Feeding system 30
3.4.1 Pressure Fed System 30
3.4.2 Other Feeding Systems 32
3.5 Design section Design of Injector 34
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Chapter IV SUB-SYSTEMS
4.1 Cooling system 42
4.2 Ignition system 44
4.2.1 Pyrotechnic Ignition System 44
4.2.2 Hypergolic Ignition System 44
4.2.3 Spark Igniters and Augmented Spark Igniters 45
4.3 Control and Measuring Systems 45
4.3.1 Gaseous Nitrogen Regulator 45
4.3.2 Gaseous Oxygen Regulator 46
4.3.3 Propellant Control Valves 46
4.3.4 Check Valves 46
4.3.5 Pressure Gauges 46
4.3.6 Temperature Indicator 46
4.3.7 Thrust Indicator 47
4.4 Design section Design of cooling system 47
4.5 Design section Sketch of the assembly 49
Chapter V MANUFACTURING AND ASSEMBLY
5.1 A Brief Description of Machine Tools 555.2 Machining process 56
5.2.1 Combustion Chamber 56
5.2.2 The Nozzle 59
5.2.3 The Injector 61
5.2.4 Engine Mount and Test Stand 62
5.3 Electroplating 63
5.4 Assembly 64
5.4.1 Engine 645.4.2 Engine mount and Screw rod 66
5.4.2 Measuring systems 66
Chapter VI TESTING
6.1 Safety Precautions 68
6.2 Engine Operation 69
CONCLUSION 70
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LIST OF TABLES
Sl.No. Section Title Page no.1 1.3 Some standards for amateur rocket designers 52 2.5 Design mach numbers corresponding to various chamber
pressures
18
3 2.5 Nozzle coordinates 22
4 3.2 Combustion chamber coordinates 27
LIST OF FIGURES
Sl.No. Section Title Page no.
1 1.1 Balloon motion 3
2 1.1 Schematic of a Liquid propellant Rocket engine 3
3 2.2 Stream line motion in a fluid particle 10
4 2.3 Nozzle 11
5 2.3 Area-Mach Relation 14
6 2.4 Convergent nozzle 16
7 2.4 De-Lavel nozzle 17
8 2.4 Normal Bell Nozzle and linear Aerospike 189 2.5 Aerospike main window 19
10 2.5 Aerospike selection of propellants 20
11 2.5 Nozzle contour 21
12 2.5 Mach. No. distribution 21
12 2.5 Pressure distribution 21
13 2.5 Temperature distribution 22
14 2.5 Density distribution 22
15 3.2 2D view of the designed engine 28
16 3.2 3D view of the designed combustion chamber 2817 3.2 3D view of the designed nozzle 28
18 3.2 3D view of the assembled engine 29
19 3.3 Impinging injector and spray injector 30
20 3.4 Schematic of a pressure fed system 31
21 3.4 Schematic of a expander feeding system 33
22 3.4 Schematic of a gas generator feeding system 33
23 3.4 Schematic of a staged combustion feeding system 33
24 3.5 2D view of fuel injector plate 34
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25 3.5 3D view of fuel injector plate 35
26 3.5 2D view of fuel injector housing 35
27 3.5 3D view of fuel injector housing 3528 3.5 3D view of the oxygen injector 37
29 3.5 3D view of the oxidizer injector 37
30 3.5 2D view of the oxidizer injector 38
31 3.5 2D view of the oxidizer housing 39
32 3.5 3D view of the oxidizer housing 39
33 3.5 3D view of the full injector 40
34 3.5 3D view of the full injector 40
35 4.1 Double wall dump cooling system 43
36 4.1 Tubular dump cooling system 4337 4.4 Cooling system passage for combustion chamber 47
38 4.4 3D view of the cooling system passage 48
39 4.4 Arc length of the helix 48
40 4.5 Thermocouple provisions in the combustion chamber 50
41 4.5 Cr/Al thermocouple with indicator 50
42 4.5 Pressure tapping provisions 51
43 4.5 Load cell with indicator 52
44 4.5 Block diagram of hypothetical engine assembly 53
45 5.1Schematic of lathe and milling machine operations
5646 5.2 Raw material for combustion chamber 56
47 5.2 Raw material for nozzle 56
48 5.2 Combustion chamber being bored in conventional lathe 57
49 5.2 Combustion chamber being machined in VTL 57
50 5.2 Finished combustion chamber 58
51 5.2 Finished combustion chamber with provisions for coolingsystem and thermocouples
59
52 5.2 Pre-machined nozzle material 59
53 5.2 Simulated views of the machining by EdgeCAM 2009 R1 60
54 5.2 Simulated views of the machining by EdgeCAM 2009 R1 6055 5.2 Simulated views of the machining by EdgeCAM 2009 R1 60
56 5.2 Actual manufactured rocket Engine Nozzle 61
57 5.2 Assembled injector 62
58 5.2 Mounted load cell on test stand 62
59 5.2 Test stand 63
60 5.2 Engine mount with screw rod 63
61 5.3 Views of Electroplated components 64
62 5.3 Views of Electroplated components 64
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63 5.4 Fastener provisions for combustion chamber and nozzle 65
64 5.4 Assembled Combustion chamber and Nozzle 65
65 5.4 Assembled Thermocouple 6666 5.4 Assembled Load cell 66
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Chapter I - INTRODUCTION
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INTRODUCTION
The power plant is the most essential part of any mechanical devices or instruments or systems.
The prime goal of any power plant is to provide enough energy to drive that system. All it does is it
converts energy from one form to another form. Power plants can be classified into many sub categories.
There is no constant rule for classification of power plants, depending upon the requirements,
manufacturability, type of energy being converted, etc., these can be sub divided into many categories.
Rocket engines are a kind of power plants which converts chemical energy into kinetic energy.
Combustible chemicals are combusted in the combustion chamber of a rocket engine and are allowed to
expand through the nozzle. These chemical rocket engines can be further classified into three types viz.,
Solid rocket motor, Liquid rocket engine and Hybrid rocket motor. In Solid rocket motor the propellant is
a mixture of fuel and oxidizer and is stored in the solid state. Liquid rocket engines have fuel and oxidizer
both stored in liquid or gaseous form in separate tanks. The disadvantage of the solid rocket motor is that
if it was ignited it will stop only when the entire propellant gone empty until that it will burncontinuously. Hence these were used at places where high thrust is required. Examples include boosters in
rockets, missiles, etc. In liquid rocket engines the case is different the reaction can be controlled by
controlling the flow of propellants in to the combustion chamber. The disadvantage is the storage of
liquid propellants is difficult than the solid propellants and the mechanism required for the operation is
complicated. Examples include many of the rocket engines which are currently in use. The disadvantages
of these two engines have been overcome by the hybrid rocket motor. It consists of the stored fuel in solid
state and oxidizer in liquid or gaseous state which is required for the combustion. The hybrid rocket motor
have some advantages over the two of its competitors, however it is not widely used. Liquid Rocket
engines were the prime thrust producers in many of the rockets which involve satellite injection, missions
to other celestial bodies and even some of the ICBMs. In this text we will study about the design of such
engine.
1.1 INTRODUCTION TO LIQUID ROCKET ENGINES
A liquid rocket engine is a kind of rocket engine which has its own systems for producing thrust
with the available fuel and oxidizer. The fuel is the combustible matter and the oxidizer supports the
combustion, without oxidizer there will not be any combustion. In the Liquid Propellant Rocket (LPR)
engines the fuel and the oxidizer will be sprayed in minute droplets and then they are ignited in the
combustion chamber. Because of combustion the kinetic energy of gases increases this result in increase
in pressure and temperature. The highly energized gas is then made to expand through the nozzle which
produces the thrust. The operation of the rocket engine and the thrust production can be simply explained
from the following fig.
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Fig. 1.1-1 Balloon motion
The pressure in the balloon is greater than that of the atmospheric pressure. Because of high
pressure, the air inside the balloon will tend to squeeze towards where the pressure is low. When the airinside the balloon is allowed to come out from the mouth of the balloon it will come with some velocity
because of the above said pressure difference and the velocity is said as exit velocity this causes the
balloon to move forward. As the air squeezes out through the throat, a force is produced and that force
moves the balloon in the opposite direction. This force is called Thrust. This is what exactly happens in a
LPR engine (on all other chemical rockets also). According to SI system the thrust is measured in Newton
(N). One Newton is the force required to move one kilogram of an object with velocity one meter/second
in one second.
Fig. 1.1-2 Schematic of a Liquid propellant Rocket engine
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The schematic of a LPR engine is shown in Fig. 1.1-2. The fuel and oxidizer are stored in their
storage tanks and are ducted to the injector. The purpose of the injector is to mix the fuel and oxidizer and
to spray them into the combustion chamber as minute droplets. The droplets are atomized in order to
make the fuel and oxidizer mix properly and also to make the combustion complete. The oxidizer and fuel
mixture is then ignited by using an igniter. Upon ignition combustion takes place and the pressure and
temperature builds up tremendously. The pressurized gases will then starts to flow towards the low
pressure area through the throat where it is accelerated to sonic velocity and then it is expanded through
the diverging nozzle to increase the velocity at the exit. The nozzle is designed in such a way that the
nozzle exit pressure matches with the ambient pressure to obtain high burnout velocities and also to avoid
the formation of shock waves while flowing at high velocities.
1.2 FUNDAMENTALS OF ROCKET PROPULSION
Thrust is the force which propels the rocket and it mainly depends on the mass flow rate and the
exit velocity. Exit pressure of the nozzle also has slight influence in the thrust. The thrust is governed by
the basic physical laws, - The Newtons second law.
=
Where , the force is produced, is the mass and is the acceleration. Now we know that
acceleration is change in velocity with respect to time. Substituting for acceleration,
=
The product of mass and velocity is called the momentum and this equation is referred to as the
momentum equation. When this equation is applied in rocket engines the values of mass and velocity can
be either the mass or velocity of the rocket or the exit gases. Now the above equation can be re-written
as,
=
Where is the mass flow rate and ve is the effective exit velocity of the gas. This is the
generalized equation for the thrust of any rocket engine. The above equation holds good if the exit
pressure (Pe) of the gases equals the ambient pressure (P). If the exit pressureis not the same as ambient
pressure then the thrust equation is given by
= + ( )
From the basic sciences we know that the product of mass and velocity is momentum, so the first
term is the momentum thrust and the second term is the pressure thrust. The pressure thrust depends on
the expansion of the nozzle. If the expansion is too high then the exit pressure will be less than the
ambient pressure and this will cause negative pressure thrust. So a rocket nozzle has to carefully designed,
so that the exit pressure should be slightly greater than the ambient pressure. The product of thrust and the
total burning time is termed as Impulse(). Mathematically it can be expressed as = , where is the
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total thrust force over a period of time and is the burning time. Specific impulse (Isp) is the impulse per
unit weight of the propellant. Specific Impulse is one of the important factor for the selection of the
propellants and is mathematically represented as
=
=
The unit of Specific Impulse is seconds (s). Here W is the mass flow rate ( ) and this mass flow
rate is the total mass flow rate including the fuel and oxidizer. Here we introduce another variable r,
which is the oxidizer fuel ratio.
=
And this gives,
=
And
= = +
From this we can find the oxidizer and fuel flow rates for a given propellant combination for the
required mass flow rate. The oxidizer/fuel ratio r depends upon each propellant choice and some of the
widely used propellant combination and their properties are given in table 1.1[1].
1.3 SOME STANDARDS FOR AMATEUR ROCKET DESIGNERS
Sl.
No.Oxidizer Fuel
Mass
mixture
ratio
(r)
Adiabatic
Flame
Temperature
(F)
Isp
(S)
Combustion
Pressure
(psi)1 Liquid Oxygen Gasoline 2.5 5470 242 300
2 Gaseous Oxygen Gasoline 2.5 5880 281 500
3 Liquid Oxygen JP-4 2.2 5880 255 500
4 Liquid Oxygen Methyl Alcohol 1.25 5180 238 300
5 Gaseous Oxygen Methyl Alcohol 1.2 5220 248 300
6 Liquid Oxygen Liquid Hydrogen 3.5 4500 363 5007 Red fuming Nitric acid JP-4 4.1 5150 238 500
8 Fluorine Liquid Hydrogen 4.54 5084 389 500Note: All these values are obtained from the pre recorded tests and these values may slightly vary from text to text. All the above
values hold good with expansion to 14.7 psi.
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Hydrocarbon fuels are the most readily available fuels in the market and they can be handled
easily unlike any other fuels. The best choice of oxidizer for an amateur rocket builder is the gaseous
oxygen and it is the best oxidizer for most of the combustible matters. Although Liquid oxygen is the best
oxidizer for any rocket it cannot be handled by an amateur rocket builder. Regardless of the pressure
oxygen will be in its liquid form if and only if the temperature is -183C. Liquid oxygen is pale blue in
color and its storage is very difficult. Since it is a cryogenic liquid it requires double wall vessel to be
stored. So, Gaseous oxygen would be the best choice for an amateur rocket builder.
1.4 DESIGN SECTION CALCULATION OF MASS FLOW RATES
The objective of this design section is to showcase the calculations involved in the design of a
Liquid propellant rocket engine. Sample calculations corresponding to each chapter will be given at
the end of the respective chapters. The design section deals with the design of a Liquid propellantrocket engine which is capable of producing 500Kgf thrust at sea level.
Calculate the mass flow rates for a rocket engine capable of producing 500Kgf thrust at sea
level with GOX and Gasoline as propellants
The first step to be considered is the impulse equation.
=
The required thrust is 500kgf. If we choose the propellant combination as gaseous oxygen and
gasoline then the specific impulse will be 281s. Substituting the values we will get,
=
= =500
281
This is the required mass flow rate for the engine for the selected propellants. Now we know themixture ratio for the selected propellants and by this,
= + And =
= (1+)
= 1.78 / = 3.924 /
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= ()
=1.78 /
(1+2.5)
And,
= 0.508 /
= 1.272 /
= 1.12 /
= 2.804 /
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Chapter II COMPRESSIBLE FLOW
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2.1 CONTINUITY EQUATION
This is based on the Law of conservation of mass which states that matter can neither be created
nor be destroyed, Or in other words the matter (or mass) is constant. Consider a one dimensional steady
flow, The mass flow per second is AV.
density of the gas
A Area of cross section
V Velocity of the flow
According to the law, AV = constant
Differentiating the above eqn.
(
AV) = 0
() +() = 0
+ +() = 0
+ + = 0
by AV
The above equation is the continuity equation in differential form.
2.2 BERNOULLIS EQUATION FROM EULERS EQUATION OF
MOTION
This is equation of motion in which the forces due to gravity and pressure are taken into account
and is derived by considering a fluid along a streamline S. Now the forces acting on the cylindrical
element with cross section dA and length ds are
Pressure force pdA in the flow direction Pressure force + in the opposite direction Weight of the cylindrical element gdAds
+
+
= 0
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Fig. 2.2-1 Stream line motion in fluid particle
According to Newtons second law, =
+ cos =
Now acceleration is afunction with respect to displacement(s) and time(t).
= =
+
For a steady flow the acceleration is zero,i.e the rate of change of velocity is zero. So = 0
Substituting the value of a for the cylindrical element, and simplifying
cos
0
From the figure cos
,
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substituting in the above equation and simplifying
This equation is the Eulers equation of motion. Bernoullis equation is obtained by integrating
this equation and by assuming
The flow is ideal, i.e. inviscous Flow is steady, incompressible and irrotational
For an incompressible flow, is constant and therefore
2.3 ISENTROPIC RELATIONS
Isentropic flow is one in which no heat is added or taken away from the process and the process is
completely reversible. Isentropic processes merges the compressible aerodynamics with
thermodynamics. Isentropic relations can derived from the basic laws of thermodynamics.
Fig. 2.3-1 Nozzle
0
2
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Now consider the first law of thermodynamics. . Second law states that .So . is the work done by the system, and it can also be denoted as . From thiswe get .We know that Differentiating,
Now substitute in the above equation.
By perfect gas law,
Integrating between two states 1 & 2, we get,
The change in entropy over a closed system is zero. So
We know that
In the same way, Substitute in and proceed, we will get
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For further simplification consider the energy equation,
2 = +2
State 1 corresponds to combustion chamber. So the velocity at this state will be zero.
= + 2
We know that =
= +
2
= 1 +
2
But =
= 1 +
2/( 1)
=
= 1 +
( 1)2
=
= 1 +
12
= 1 + 1
2
=
=
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From this
AREA - MACH RELATION
Fig. 2.3-2 Area-Mach Relation
1
1
2
1
12
1
12
1 12
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Consider figure 2.3-2. The mass flow across any section has to be same. So
But1 = 1 + 12 2 11 and = () (since M=1)
=
=
1 + 12 22
1
By Definition,
= = + 12 1 + 12
Substituting and simplifying we get
With the help of above equations we can easily calculate all the major parameters of a nozzle
corresponding to the local Mach number. The pressure, density and temperature ratio relations help us in
finding out the local values of the respective relations. The suffixed variable represents the known value,i.e. the chamber conditions. The Area ratio relation finally gives the profile of the nozzle with respect to
the Mach number. While designing a nozzle the nozzle design Mach number is chosen and is fed up in the
formula to find the exit area. The throat area can be found out by choking the nozzle. The condition at
which the nozzle will be choked is given by
=
= 1 2 + 1 1 + 12
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2.4 THE NOZZLE
Nozzle is the most important part of a rocket engine. It is simply a duct of varying cross sectional
area and in rocket engines nozzles are used to accelerate the flow of gases. Nozzle contour is basically
obtained by the isentropic relations. The primary objective of the nozzle is to increase the kinetic energy
of the gases. Nozzles can be classified into the following types.
Convergent nozzle De-Laval Nozzle (Convergent divergent nozzle) Minimal length nozzle Plug or Aerospike
CONVERGENT NOZZLE: - A convergent nozzle is a kind of nozzle which is generally used for
increasing the velocity of the gases in the subsonic region. The principle behind this nozzle is the
continuity equation which we will be discussing later. The simple form of continuity equation is given
below. A schematic of a convergent nozzle is shown in the Fig. 2.1-1.
When the flow is incompressible, i.e., When the density is constant the equation changes to
Fig. 2.4-1 Convergent nozzle
CONVERGENT- DIVERGENT NOZZLE: - This nozzle is mainly used for accelerating the flows to
supersonic speeds. Supersonic flow can be achieved only by a C-D Nozzle or De-Lavel nozzle. The
convergent section of nozzle accelerates the flow up to transonic speeds and finally sonic speed will be
achieved at the throat when the nozzle is choked. Throat is the part where the cross-sectional area will be
at its minimum. Nozzle can be choked by further increasing the pressure ratio. Behind the throat region
the divergent section begins. Here, the pressure drops and the flow begins to accelerate. The acceleration
of the flow depends upon the expansion of the gases. The divergent section of most of the nozzles will be
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simply conical which does not yield the perfect expansion of the gases. Perfect expansion is possible only
with the bell nozzle where the contour will be parabolic. Design of such bell nozzles depends on the
method of characteristics and --M relationship. The design itself is a separate science and it involves in
the solution of higher order differential equations. Many software packages are available to solve this
criterion and software based solution will be discussed here.
Fig. 2.4-2 De-lavel nozzle
MINIMAL LENGTH NOZZLE: - This is very similar to a De-Lavel nozzle except the length of the
total nozzle will be reduced by increasing the divergent angle and also by reducing the size of the
convergent section. This reduction will be done in such a way that it should not affect the performance of
rocket engine. Most of the modern rockets have this kind of nozzle (except some experimental rockets).
Here in this text we will be designing this kind of nozzle in our example calculations.
PLUG OR AEROSPIKE: - The aerospike engine is a type of rocket engine that maintains its
aerodynamic efficiency across a wide range of altitudes through the use of an aerospike nozzle. It is a
member of the class of altitude compensating nozzle engines. A vehicle with an aerospike engine uses
2530% less fuel at low altitudes, where most missions have the greatest need for thrust. Aerospike
engines have been studied for a number of years and are the baseline engines for many single-stage-to-
orbit (SSTO) designs and were also a strong contender for the Space Shuttle main engine. However, no
engine is in commercial production. The best large-scale aerospikes are still only in testing phases. The
design of this kind of nozzle can be done with the help of software packages however we will not be
discussing about this type of nozzle.
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Fig. 2.4-3 Normal Bell Nozzle and linear Aerospike
2.5 DESIGN SECTION DESIGN OF NOZZLE
Find out the throat area required for the accounted rocket engine and hence plot the
contour of the rocket engines nozzle. Choose a design Mach number of 3.
Combustion Chamber
pressureDesign Mach
No.psi bar
100 6.894 1.95
200 13.789 2.33
300 20.684 2.55400 27.579 2.73
500 34.473 2.83
Table 2.5-1 Design Mach Numbers corresponding to various chamber pressures
The throat area can be found out by choking the nozzle and the relation is given by
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From the isentropic relations, For M=1 and =1.23 we get,
0
.559
279.5
196507.9 /
0.897 5274.36 3185.572 Value of R for combusted gases is 357 & 9.8 /.
1.78 /196507.9 / 3573185.572 1.239.81/ 2.780810
59.5
For the ease of manufacture and to ensure proper choking, round this value to 59mm. So
To find the contour we need to solve higher order differential equations which are very difficult
to solve manually. As already explained there are many software packages available to solve the higher
order differential equations in order to get the parabolic contour of the nozzles. We have chosen
AEROSPIKE 2.6 given by aerorocket.com. The following screen shots will best explain the datas to be
fed into the software and the output given by it.
Fig. 2.5-1 Aerospike main window
59
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The above image is the main window of the software. As indicated in the above figure feed the
pressure ratio as 34.01 in the provided area. Select the propellant combination as shown in the following
figure. The fuel we are using is gasoline, but this option is not available in this software. So select RP-1
(Rocket Propellant 1) and then change the value of as 1.23.
Fig. 2.5-2 Aerospike selection of the propellants
After this, click on the button over Minimum Length Nozzle. A new window will appear and in
that window change the value of to 1.23, design Mach No. to 3 and Throat diameter to 0.059 m as
shown in the following fig.
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Fig. 2.5-3 Nozzle contour
By clicking the plot button various distributions over the nozzle can be figured out.
Fig. 2.5-4 Mach No. distribution Fig. 2.5-5 Pressure distribution
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Fig. 2.5-6 Temperature distribution Fig. 2.5-7 Density distribution
Now click on the Send MLN data to printer button and get the coordinates as shown below.
Table 2.5-2 Nozzle coordinates.
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So here we have found out the throat area and hence the contour of the nozzle has been generated.
All the obtained values can be checked for consistency by manual calculations. For the ease of
manufacture the nozzle and combustion chamber can be hypothetically divided into two sections. These
two sections can be brought together and fixed up by flanges which are fitted on them. The thickness is
the next problem which we will be discussing in the subsequent sections.
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Chapter III THE COMBUSTION CHAMBER
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3.1 COMBUSTION CHAMBER
Combustion chamber or combustor is the place where all the fuel and oxidizer gets burned up and
the total pressure, temperature and the energy gets tremendously increased. The purpose of combustion
chamber is to provide the required pressure at the throat of the nozzle so as to obtain the supersonic flow
at the exit. Combustors are designed to contain and control the burning fuel-oxidizer mixture. The
combustor normally consists of two components: an outer casing that acts as a high pressure container
and the fuel injection system. Here we will be discussing about implementation and design of these two
components.
The first one is the high pressure container i.e., the combustion chamber. Combustion chamber
volume is the major parameter in the design and is obtained by assuming a characteristic length which is
required for complete combustion. Generally the characteristic length (L*) is assumed to be 20 inches
(0.508 m). Then the volume is calculated by using the following relation. The convergent section volumeis also included in this.
The radius of the combustion chamber is generally taken 2-3 times the radius of the throat. With
this radius the chamber cross sectional area (Ach) is found out and the actual length of the combustion
chamber (Lch) is found out by the following method. The decimal 1.15 in the denominator is to neglect the
length of the convergent section. This means 10% of the total volume is allotted for the convergent
section.
For most of the engines the volume of the convergent section is generally 15 100 of the totalvolume. So,
= 1.15
=
1.15
Here we have found the length and cross sectional area of the combustion chamber, and now the
length of the convergent section is found out by volume of the truncated cone (frustum) formula.
=
3
+ +
Where, Vcon is the convergent section volume. (15% of the total volume)
Rch is the radius of the combustion chamber
R* is the radius of the throat.
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With these datas we can easily plot the combustion chamber. Now since this is a cylinder we can
easily calculate the thickness required by thin shell formula,
2
Here is the pressure inside the cylinder, is the working stress and is the diameter of the
cylinder or the combustion chamber. Knowing all these values the thickness of the chamber can be
easily calculated. For calculating the thickness we need to know the working stress of the material.
Although material selection lies in the manufacturing phase, it is must here for estimation of thickness of
the components and so to plot the contour of the components. Modern rocket engines are fully made of
composite materials, since composites have extremely high strength/weight ratio and high strength than
any other materials. But machining the composites is extremely hard and cannot be machined in a similar
way that we machine the metals. Composite materials are set of laminates binded together by a resin or
any other similar adhesive material. Each laminate consists of several laminas. Every lamina is formed by
fibers and matrix. A fiber is simply a synthetic thread which is laid down in a net like pattern and a matrix
is formed by separate process. So if we need to manufacture a whole engine in composite we need to lay
down the laminas exactly to the engines profile and then the whole of the engine has to be developed in a
similar process. The description of this technique lies beyond the scope of this text and hence machinable
materials will be chosen and will be discussed here.
Depending upon the cost available and manufacturing facility available the materials has to be
chosen. Most of the high temperature materials used in Aerospace industry costs more and so an amateur
designer may not be able to spare. Some of the high temperature materials are Hastalloy series, Nickel
alloys, Chromium alloys, Molybdenum alloys, titanium alloys, Ni-Cr alloys, Graphite, Vanadium steel,
Silicon added alloys. It should be noted that not all these may be readily available in market and the most
easily available metals are Stainless steel, copper and mild steel only.
3.2 DESIGN SECTION DESIGN OF COMBUSTION CHAMBER
Design a thruster for the accounted rocket engine, provided with mild steel EN8. Hence
estimate the thickness of the combustion chamber and assume it as the uniform thickness
throughout the nozzle. Plot the contour of the combustion chamber and the nozzle.
=
= 2.7808 10
0.508
= 1.4126 10
Now the radius of the combustion chamber is taken as 2 times the radius of throat. Therefore = 2
29.5 10
0.059 . So the area is 0.059 = 0.0109358 .
=
1.15 =
1.4126 10
1.15 0.0109358 = 112
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This is the total length of the combustion chamber. For the ease of manufacture the tolerance
limit can be made up to 5 ( 5) and the diameter of the combustion chamber can be
round up to
0.06. Now the length of the convergent section is found out by the following relation. The
value of is 15% of the total volume i.e. 1.41 103 0.15 = 2.1 1043
=
3
+ + =
2.110
3 (0.06 + 0.02975 + 0.06 0.02975)
= 31.4
So the coordinates for the combustion chamber is given by
X (mm) Y (mm)
0 0
0 120
85.6 120
117 89.5117 30.5
85.6 0
Table 3.2-1 Combustion chamber coordinates
Now we need to calculate the thickness of the combustion chamber. Since the material to be
chosen is EN8 (Emergency Number 8), the yield strength of EN8 is 250Mpa. Considering a factor of
safety as 2, the working stress () can be calculated as
= 125 . So by thin shell formula,
= 2
= 3.447379 120 2 125
2
Now for safety purposes make this thickness to 5mm for the actual engine. For the hypothetical
implementation of cooling system (which we will be discussing later) fix the thickness as 10 mm and
make it uniform throughout the engine.
It should be noted that the flanges has to be designed in such a way that it should be capable of
holding the bolts and nuts. Though the diameter of the hole for fastening is assumed as 10 mm the total
diameter of the head and the nut will be more. So care should be taken while designing the flanges. The
following figures best explains the design of combustion chamber and nozzle.
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Fig. 3.2-1 2D view of the designed engine (All dimensions are in mm)
Fig. 3.2-2 & 3 3D view of the designed combustion chamber (left) and Nozzle
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Fig. 3.2-4 3D view of the assembled engine
3.3 THE INJECTOR AND ITS TYPESThough small in size the injector is the major part in a rocket engine. It mixes the fuel and
oxidizer in proper ratio and makes the mixture to flow in the proper direction.The injectors in a chemical
rocket motor are the key in determining the efficiency of the reactions within the combustion chamber,
ultimately affecting the performance of the motor. Critical to achieving good performance is the
atomization process, whereby the propellant and oxidizer are transformed into small droplets; in essence
the size of these drops determines the mixing process and evaporation rates, which have a profound
influence on the combustion reactions. The basic function of the injector in a bipropellant liquid rocket is
to atomize and mix the fuel with the oxidizer to produce efficient and stable combustion that will produce
the required thrust without endangering hardware durability. Currently, most bipropellant rockets and
hybrid rockets use small orifices in the injector plate, which takes the form of a perforated disk at the headof the combustion chamber.
There are generally two types of injectors. One is impinging type and the other is spray type.
Spray type injectors are the ones which are normally used in automobiles. Impinging type injector is the
one which is specifically used in rocket engines. However if the engine is a small one, spray type injector
can also be used. But for big engines impinging type has to be used. The schematic of both the injectors
were shown in the following fig. The impinging stream type injector is very difficult to manufacture but
the design can slightly be altered for the ease of manufacture.
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Fig. 3.3-1 Impinging injector and spray injector
The design of such impinging injector plate itself is a separate science and hence a basic injector
which is capable of inducing the ignition will be explained here in this text.
3.4 FEEDING SYSTEM
The feeding system is the major system which is involved in the design of injector plate, since the
injector design mainly depends upon the mass flow rates of the fuel and oxidizer. There are four different
types of feeding system which are currently in use. They are pressure fed system, expander cycle, gas
generator cycle and staged combustion cycle. In these systems staged combustion system is a
combination of expander system and gas generator system.
3.4.1 Pressure Fed System
If the engine designed is a small one, then the mass flow rates will also be less. If the mass flowrates were small then the fuel flow can be made possible by a reservoir fitted at an elevated area. Through
a small orifice (say 1 inch dia) the fuel can be ducted to the injector and the velocity of the flow at the end
of the duct can be found by the formula 2. Where H is the total height of water level from thelower end of the duct where the injector is connected. The total discharge can be found out by =
. Where is the discharge coefficient and is approximately equal to 0.6. If the mass flow rates are
higher, then a separate pump has to be used or the reservoir has to be pressurized. Pressurizing the
reservoir is one of the complicated methods both for manufacturing and handling. However if one needs a
pressurized fuel container for the required mass flow rate, then the required pressure can be found out by
the relation,
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(/)
(/
) (/
)
() (/)
2 (/
)
Here the cross sectional area of the container can be anything. is the velocity of the flow at the
orifice at the bottom of the container, considered the orifice is at the bottom. H is the total height of the
water level from the bottom. If the orifice is not at the bottom, then H should be taken as height of water
level from the centre of the orifice. is the density of the fuel and g is acceleration due to gravity. is the
required pressure over the reservoir to maintain the velocity.
After finding the velocity the discharge can be calculated by = . Where is the cross
sectional area of the orifice. With these relations proper Height of fuel and the required pressure in order
to deliver the required mass flow rate can be calculated. The following figure simply explains a pressure
fed system.
Fig. 3.4.1-1 Schematic of a pressure fed system
This method holds good if one doesnt want to take care of pumps, etc. The alternate method is to
use pumps driven by external source or any other three different systems which will be briefed in the
subsequent section. This method is very suitable for any amateur rocket builder and doesnt involve any
complex design calculations and experiments. The implementation of this system will be explained here
in the subsequent design section and it can be easily understood there.
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3.4.2 Other Feeding Systems
EXPANDER CYCLE
Heat in the cooling jacket of the main combustion chamber serves to vaporize the fuel. The fuel
vapor is then passed through the turbine and injected into the main chamber to burn with the oxidizer.
This cycle works with fuels such as hydrogen or methane, which have a low boiling point and can be
vaporized easily. As with the staged combustion cycle, all of the propellants are burned at the optimal
mixture ratio in the main chamber, and typically no flow is dumped overboard; however, the heat transfer
to the fuel limits the power available to the turbine, making this cycle appropriate for small to midsize
engines. A variation of the system is the open, or bleed, expander cycle, which uses only a portion of the
fuel to drive the turbine. In this variation, the turbine exhaust is dumped overboard to ambient pressure to
increase the turbine pressure ratio and power output. This can achieve higher chamber pressures than the
closed expander cycle although at lower efficiency because of the overboard flow.
GAS GENERATOR CYCLE
The gas-generator cycle, also called open cycle, taps off a small amount of fuel and oxidizer from
the main flow (typically 3 to 7 percent) to feed a burner called a gas generator. The hot gas from this
generator passes through a turbine to generate power for the pumps that send propellants to the
combustion chamber. The hot gas is then either dumped overboard or sent into the main nozzle
downstream. Increasing the flow of propellants into the gas generator increases the speed of the turbine,
which increases the flow of propellants into the main combustion chamber, and hence, the amount of
thrust produced. The gas generator must burn propellants at a less-than-optimal mixture ratio to keep the
temperature low for the turbine blades. Thus, the cycle is appropriate for moderate power requirements
but not high-power systems, which would have to divert a large portion of the main flow to the less
efficient gas-generator flow.
STAGED COMBUSTION CYCLE
In a staged combustion cycle, also called closed cycle, the propellants are burned in stages. Like
the gas-generator cycle, this cycle also has a burner, called a pre-burner, to generate gas for a turbine. Thepre-burner taps off, and burns a small amount of one propellant and a large amount of the other,
producing an oxidizer-rich or fuel-rich hot gas mixture that is mostly unburned vaporized propellant. This
hot gas is then passed through the turbine, injected into the main chamber, and burned again with the
remaining propellants. The advantage over the gas-generator cycle is that all of the propellants are burned
at the optimal mixture ratio in the main chamber and no flow is dumped overboard. The staged
combustion cycle is often used for high-power applications. The higher the chamber pressure, the smaller
and lighter the engine can be to produce the same thrust. Development cost for this cycle is higher
because the high pressures complicate the development process. Further disadvantages are harsh turbine
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conditions, high temperature piping required to carry hot gases and a very complicated feedback and
control design.
These three methods uses extreme technology and it cannot be easily employed in the first design
itself. So the best choice for amateur rocket developers is to run the pump through external power source.
Fig. 3.4.2-1, 2 & 3 Schematic of expander, gas generator and staged combustion feeding systems
Oxidizer flow is another important thing which is to be accounted. If the oxidizer is in liquid form
then the above mentioned method (i.e. for fuel) can be used. Note that if the oxidizer is cryogenic (E.g.
Liquid oxygen) then normal pumps cannot be used to pump this. Cryogenic pumps are available and
those pumps must be used for pumping LOX. Pressurization system can also be used but the technology
exists beyond the knowledge of amateur rocket developers. Any other oxidizers which are in liquid form
under STP can be handled easily and pressurization system can be used.
For amateur rocket developers commercially available gaseous oxygen which is available in
portable cylinders will help the most and gaseous oxygen based example will be discussed in our design
section. Even the outlet dia of any portable cylinder is 1 inch (25.4 mm,) the throat dia is only 3mm. This
is the major value while taking into account. The pressure inside a portable gaseous oxygen cylinder is
150 Kg/cm2.
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3.5 DESIGN SECTION DESIGN OF INJECTOR
Find the total fuel discharge and hence estimate the power required for a pump to feed fuel
and also find the pressure at which the oxygen from the cylinder to be released provided with two
oxygen cylinders. Also design an impinging injector that best suits the application and should be
capable of being manufactured easily.
The fuel we have chosen in gasoline. The density of gasoline is 44.5 lb/ft3. The total mass flow
of fuel is 1.12 lb/s. We know that /.Volume flow i.e. .
. 0.0251 / = 2565 . So
this flow can be made possible by a commercially available pump with discharge of approximately 3000
LPH. In case if someone needs to pressurize his container then the above mentioned method can be used.
The diameter of the outlet of the pump (probably 0.5 HP) will be 1 inch (25.4 mm). With a collar
or reducer, reduce the diameter to 0.75 inch, so that the duct will have a uniform diameter of 0.75 inch.
When the injector plate is reached the diameter will increase to spread the fuel over large number of tiny
holes. In this case the total cross sectional area for the fuel flow is taken as the area of a half inch duct
(126.676 mm2) so as to increase the velocity at the end. No.60 drill (Dia 1.016mm) is chosen and the area
of each drill is given by 0.8107 mm2. So the number of holes will 126.676/0.81 = 156 .Out of
these 40 holes can be made with Pitch Center Diameter (PCD) of 38mm and 52 holes can be made with a
PCD of 58 mm and 64 holes can be made with a PCD of 78 mm. The following figure best explains the
criteria.
Fig. 3.5-1 2D view of fuel injector plate(All dimensions are in mm)
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Fig. 3.5-2 3D view of fuel injector plate
Now we need to design a housing assembly to cover this plate and so as to allow the injector plate
to get connected to pipe lines. This can be done in a simple way as shown in the following figure.
Fig. 3.5-3 & 4 2D and 3D views of fuel injector housing (All dimensions are in mm)
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So here we have designed the perforated plate and its connection assembly to pump. Now we
need to design the injector plate for the oxidizer which should impinge on the fuel flow to make the
perfect mixture.
Density of gaseous oxygen at STP is 0.08165 lb/ft3
(1.308 Kg/m3). But the gas is stored at 150
Kg/cm2 (1.471107 N/m2) in portable oxygen cylinders which are commercially available. So its density
at that pressure can be calculated as follows.
1.47110287288 =177.845 / = 11.1025 /The required oxidizer mass flow rate is 1.272 Kg/s (2.804 lb/s). If we open the oxidizer cylinder
the density varies with respect to the pressure. So let us use a pressure regulator to regulate the overall
pressure to 100Kg/cm2
. At this pressure the density of oxygen will be 118.64 Kg/m3
(7.406 lb/ft3
). Nowwe need to calculate the velocity of the flow from this pressure (100 Kg/cm2 or 1422.334 psi) to the
combustion chamber pressure (35.153 Kg/cm2
or 500 psi). For this we need the density of oxygen at 500
psi and is found to be 41.707 Kg/m3 or 2.603 lb/ft3.
Applying Bernoullis principle,
+ 12 = + 12 Here state 1 corresponds to the regulator location and state 2 corresponds to the combustion
chamber location. Since state 1 is a reservoir = 0. So, = 2( ) = 2(1000000 /
351530 /)41.707 / =176.3 /Now we need to choose the diameter of the ducts, Let us choose a flexible hose of diameter 0.5
inch (12.7 mm). The total mass flow rate which can be made possible by a single cylinder is givenby = . Near the injector plate the density is 41.707 Kg/m3. The cross sectional area is (12.710) =1.266710 and the velocity is 176.3 m/s. Therefore the mass flow rate is0.9314Kg/s. For 2 cylinders 1.8628 Kg/s
But we need a total mass flow rate of 1.272 Kg/s. For 1 cylinder the required mass flow is 0.636
Kg/s. and A will not change. So the velocity has to be reduced to satisfy the required conditions. For
this mass flow rate, the required velocity is
= = 0.636 /41.707 / 1.266710 =120.38 /
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After calculating the velocity, apply the Bernoullis principle in the reverse way to calculate the required
pressure.
120.38 / 2 351530 /41.707 /
So even if we use a 100Kg/cm2 regulator we need to regulate the pressure to only 66 Kg/cm2 in
both the cylinders to achieve the required mass flow rate.
Since we will be using 2 cylinders there will be 2 0.5 inch hoses for ducting the pressurized
oxygen to the injector. Two inlets on the oxidizer injector plate will do the job, but let us make it as four
for proper distribution. This increase in area will be later reduced to compensate the pressure loss. After
that section we have to distribute the oxidizer over a large area for injecting them through the inclined
perforated ring. This large area will reduce the pressure for a little amount and then the pressure get builds
up before the injection into the combustion chamber if we reduce the flow area. Now the flow area over
the inclined ring is calculated as follows. The flow area around the housing of the oxygen inlet is given
by 117 100 3000 . Now let us reduce this area to half of this value to increase thevelocity and pressure of the injection. So the area over the inclined ring for the injection is 1500 mm2.
This area has to be equally divided on the inclined ring as holes. For manufacturing feasibility let us
choose the number of holes as 30 and these yields the dia of each hole as 30.
It should be noted that the injector which is designed here is based on the convenience of the
authors. The objective is to distribute the flow into the combustion chamber through perforated discs or
rings without any loss. The injector which is designed here may be complex to understand and any new
kind of design can be adopted depending on the manufacturing feasibility and the flow distributions. For
convenience the 2D and 3D views of the designed injector is shown in the upcoming figures.
Fig. 3.5-5 & 6 3D view of the oxidizer injector
65.3725 /
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Fig. 3.5-7 2D view of the oxidizer injector (All dimensions are in mm)
Now as similar to the fuel housing the following figure shows the oxidizer housing and this
completes the design of the fuel injector.
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Fig. 3.5-8 2D view of the oxidizer housing (All dimensions are in mm)
Fig. 3.5-9 3D view of the oxidizer housing
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As stated earlier the injector can be designed in any way so that it should deliver the total mass
flow rates of the fuel and oxidizer to the combustion chamber in a well mixed manner. If the injector
doesnt mixes the fuel and oxidizer in a good way then combustion instability will occur which may result
in disaster. The assembled view of the designed injector is shown below.
Fig. 3.5-10 3D view of the full injector
Fig. 3.5-11 3D view of the full injector
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Chapter IV SUB-SYSTEMS
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4.1 COOLING SYSTEM
The rocket engine is the place where extreme temperatures were reached. When the flame
temperatures reach 3500K, then the state of the matter will be plasma which is nearly equal to the plasma
state of the sun where the surface temperature is 5800K. Anyhow this flame temperature will not
immediately melt and vaporize the combustion chamber/nozzle material. Most of the high temperature
metals will have their melting point around 1800C. Some non metals like silica based compounds;
composites, etc will have a higher melting point in the order of 3000C. To effectively use this advantage
of composites modern rocket engines were fully made up of composites. Anyhow even if the engines
were fully made up of composites cooling is essential to improve the overall performance of the engine.
Because of high combustion temperatures and high heat transfer rates from the hot gases to the
combustion chamber wall thrust chamber cooling becomes a major consideration. For short duration un-
cooled chamber walls can be used. In this case the heat has to be absorbed by the sufficiently heavycombustion chamber which acts as the heat sink. For most long duration applications a steady state
cooling system has to be employed. Some of the cooling systems used in rocket engines are as follows.
1. Regenerative cooling2. Dump cooling3. Film cooling4. Transpiration cooling5. Radiation cooling6. Ablative cooling
REGENERATIVE COOLING: is the most widely used method of cooling a thrust chamber and is
accomplished by flowing high-velocity coolant over the back side of the chamber hot gas wall to
convectively cool the hot gas liner. The coolant with the heat input from cooling the liner is then
discharged into the injector and utilized as a propellant. For the fuel/oxidizer to flow around the
combustion chamber/nozzle, the path has to be formed. The path can be formed either longitudinally or
spirally. Spiral path has many advantages than longitudinal paths. For these either double walled
combustion chamber or single wall with tubular winding can be done over it. Most of the engines with
regenerative cooling system were employed with spiral tubular winding type of ducts for the cooling
purposes. Since it involves lots of thermodynamic calculations these systems cannot be easily designed in
the first design itself.
DUMP COOLING: It is similar to regenerative cooling because the coolant flows through smallpassages over the sides of the thrust chamber wall. The difference is that after cooling the thrust chamber,
the coolant is discharged overboard through openings at the aft end of the divergent nozzle. Here we will
be using such kind of cooling system in our design section. As similar to the flow passages around the
combustion chamber/nozzle in the regenerative cooling system, the passage has to be made similar to that
and is shown in the following figure.
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Fig. 4.1-1 Double wall dump cooling system (longitudinal and spiral)
Fig. 4.1-2 Tubular dump cooling system (longitudinal and spiral)
The heat transfer from the hot combustion chamber gases to the chamber wall is by convection.
The amount of heat transfer through conduction is very small and the radiation heat transfer will be
usually less than 25% of the total heat transfer. Now the amount of heat which can be transferred to the
coolant is given by
( )
Where, is the total heat transferred (Btu/s)
is average heat transfer rate of chamber (Btu/in2 S)
A is heat transfer area
Ww is mass flow rate of coolant (lb/s)
Cp is specific heat of coolant (Btu/lb F)
T is the temperature of coolant leaving the jacket (F)
Ti is the temperature of coolant entering into the jacket (F)
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FILM COOLING: It provides protection from excessive heat by introducing a thin film of coolant or
propellant through orifices around the injector periphery or through manifold orifices in the chamber wall
near the injector or chamber throat region. This method is typically used in high heat flux regions and in
combination with regenerative cooling.
TRANSPIRATION COOLING provides coolant (either gaseous or liquid propellant) through a porous
chamber wall at a rate sufficient to maintain the chamber hot gas wall to the desired temperature. The
technique is really a special case of film cooling.
RADIATION COOLING: In this process heat is radiated away from the surface of the outer combustion
chamber wall, when the chamber is red hot. This requires high temperature wall material (possibly
refractory material).
ABLATIVE COOLING: In this process a sacrifice of combustion chamber gas side wall material is
made by melting and subsequently vaporizing it to dissipate heat. As a result relatively cool gases flowover the wall surface thus creating a cooler boundary layer assisting the cooling process. An ablative
material is a fiber-reinforced resin that pyrolyzes endothermally within the chamber wall releasing a cool
gas over the inner surface of the combustion chamber. The best ablative coating which is especially used
in rocket engines is Dow Corning DC 93-104. If this is not available commercially then MOLYKOTE P-
37 Anti-seize paste will nearly be a good substitute for DC 93-104.
4.2 IGNITION SYSTEM
Igniters were the devices which release heat and thus initiating the combustion. Igniters have theirpower from an outside source or from a limited amount of energy stored as solid propellants within
themselves. After the initiating the combustion in the main combustion chamber, the igniter doesnt
involves in further ignitions. The ignition can be given to the engine based on any technique that is safer
and efficient. Some of the widely used ignition systems were discussed below.
4.2.1 Pyrotechnic Ignition System
Pyrotechnic igniters are simply fire crackers with low burning rates. They consist of stored solid
propellants and usually their burning time will be less. The implementation is very simple so that it can be
mounted anywhere on the thrust chamber and usually they are screwed at the center of the injector plate.
To prevent hard start (sudden combustion of more fuel due to delayed ignition) and to ensure good
starting more than one pyrotechnic igniter can be placed in the combustion chamber. It should be noted
that pyrotechnic igniters were mainly suitable for small engines and depending upon the size of the main
combustion chamber the size of the igniter has to be increased which is not practically possible. Also it
has many disadvantages when cryogenic propellants were used.
4.2.2 Hypergolic Ignition System
The term hypergolic defines a bipropellant combination (Hydrazine hydrate N2H4.H2O and
Hydrogen peroxide H2O2 in 1:4 ratios) which spontaneously gets fired when they are in contact. Most of
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the American engines used hypergolic propellants for the main combustion, since it doesnt require a
separate ignition system. Also the possibility of hard start is zero when hypergolic fuels were used. Now
as igniter a small impinging is nozzle is designed and a little amount of these hypergolic propellants were
fed through small pipelines. The flow of these hypergolic propellants can be turned on whenever ignition
is required. This system is mainly suitable for large engines and the storage of the hypergolic propellants
is a challenging task.
4.2.3 Spark Igniters and Augmented Spark Igniters
Spark igniters are electrical devices which uses a rotating magnet and capacitors to produce the
spark. The distance between the leads should be around 1 mm, so that when electrical power is switched
ON an arc with voltage around 10000V jumps between the leads. Augmented spark igniters are a
developed version of spark igniters. These igniters use a spark igniter and a small gas generator to
produce the flame. A little amount of fuel and oxidizer are sprayed in the gas generator and a spark igniteris used to ignite the mixture. When ignited the flame from the gas generator is used to ignite the main
combustion chamber. This is the most widely used technique for the ignition of the modern rocket
engines.
4.3 CONTROL AND MEASURING SYSTEMS
As the name indicates these systems maintains the degree of accuracy and safety. Any engine
should not be made to run without a proper control system in order to prevent disaster. Since a rocket
engine is extremely advanced application of the technology it should be handled and controlled safely.Some of the major control and measuring systems which are required for the operation of a rocket engine
test bed are described as follows.
4.3.1 Gaseous Nitrogen Regulator
The purpose of a regulator is to maintain a constant pressure on the downstream side of the
regulator as the pressure in the gas cylinder on the upstream side decreases. A good quality regulator will
maintain the downstream pressure quite accurately over a range of gas flow rates as long as the upstream
cylinder pressure does not decrease so as to become too close to the downstream pressure. Thus, all the
gas in the cylinder is not usable since some excess pressure (hence, gas) is required to drive the gas
through, and maintain control of the regulator. The flow rate of nitrogen gas required for the fuel from the
tank is relatively small and could be handled by a regular gaseous oxygen welding regulator equipped
with nitrogen cylinder fittings. However, most welding regulators do not permit adjustment to the high
downstream pressure required for rocket engine operation since welding operation requires only less
pressure in the downstream. So some special high pressure regulators have to be chosen for this purpose.
Since we are not using the pressure fed system for fuel, this regulator will not be discussed in the design
section.
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4.3.2 Gaseous Oxygen Regulator
This is for controlling the mass flow rate of the oxidizer into the engine. The discussion of
regulators for gaseous nitrogen service applies to gaseous oxygen also, except that the regulator should be
especially cleaned for oxygen service. Regulator manufacturers should be consulted for recommendations
on seat materials for use with gaseous oxygen in their regulators. Special fittings for attaching the
regulator to the oxygen cylinder are available from the sources supplying nitrogen cylinder fittings. These
sources can also supply cylinder manifold kits so that two or more oxygen cylinders can be used
simultaneously to achieve long engine run durations. The high downstream pressure regulators are also
required here for controlling the mass flow rate of the oxidizer into the combustion chamber.
4.3.3 Propellant Control Valves
These valves were used to control the flow rate into the combustion chamber. For controlling themass flow of fuel a gate valve can be used. But incase if the mass flow has to be immediately stopped
then gate valve will not help in that situation. So a ball valve has to be fitted on the fluid line after the gate
valve. When oxidizer is taken into account a ball valve will nearly control all the problems, since the
oxidizer is already regulated. Also the use of such gate valves in gases gives poor results and hence ball
valve must be used for turning OFF and ON operations.
4.3.4 Check Valves
Check valves permit fluid flow in one direction only. They are widely used in the aircraft and
hydraulic industry and are manufactured by many companies. Depending upon the size of the fluid line
the check valve has to be chosen. Check valves should be thoroughly cleaned prior to use and tested toinsure that the check is working properly.
4.3.5 Pressure Gauges
Pressure gauges are must for measuring the upstream and downstream pressures from the
regulators. Also performance parameters like combustion chamber pressure, throat pressure and nozzle
exit pressure have also to be measured. All these can be made possible by a commercially available
Bourdon tube pressure gauge or any other digital pressure transducers for more accuracy. But care should
be taken while measuring the performance parameters since the temperature of the hot gases is very and
that may cause damage to the indicator itself.
4.3.6 Temperature Indicator
As a part of performance parameters, temperatures at combustion chamber, throat and nozzle can
be measured with the help of a commercially available thermocouple. Depending upon the temperature
range the thermocouple material has to be chosen. Cr/Al alloy thermocouple can be used to measure
temperatures up to 1200C and Pt/Pt/Cr alloy thermocouple can be used to measure temperatures up to
1600C.
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4.3.7 Thrust Indicator
Thrust is the ability of the rocket engine and is the major performance parameter which is to be
measured. This can be measured with the help of a load cell. Nowadays load cells are widely used in
many places and depending upon the calculated thrust level the load cell has to be chosen and calibrated.
The mounting of the load cell may vary depending upon the manufacturer and possibly a truck weighing
load cell with low capacity is recommended for measuring the thrust.
4.4 DESIGN SECTION DESIGN OF COOLING SYSTEM
With water as coolant design a dump type cooling system for the combustion chamber alone
and hence estimate the mass flow rate of the coolant.
( )
Where, is the total heat transferred from the combustion chamber to the coolant. Now, the total heat
transferred is calculated as
11.3 /
Now let us design a cooling passage as shown in the following figure.
Fig. 4.4-1 Cooling system passage for combustion chamber (All dimensions are in mm)
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Fig. 4.4-2 3D view of the cooling system passage
This is very similar to a thread of pitch 20mm. The outer radius of combustion chamber is 70mm.
Then the helical path is formed over a depth of 5mm. So when considered as a circle the radius is 65mm.
Now consider the following figure to find the length in 1 turn.
Pitch = 20mm
Circular circumference = 2r
= 2**65 = 408.2 mm
Helical circumference
Fig. 4.4-2 Arc length of the helix
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tan 20408.2 =2.805
Therefore the length of the helical circumference is found out by
= 20 sin2.805 =
200.04893 =408.689
This is the arc length in 1 turn. The total no. of turns is 3.5. So the total arc length is 1430.4 mm.
Now the depth of the path is 5 mm and its width is 10mm. So the total surface area (A) is (1430.4*10) +
(1430*5*2) =28608 mm2. So
= 1 1 . 3 / 28608 10 = 0.323 /0.323 / =
(
)
Now 0.323 W/K is the possible heat transfer from the combustion chamber to water with the
available surface area. If water is used as coolant then, Cp for water is 4.178 KJ/Kg K, and the water inlet
temperature (Ti) will be 293K. It should be noted that the boiling point of water is 373K. So let us choose
the outlet temperature (T) as 323K and so the required mass flow can be found as,
= 0.323 /4178 (323293)= 0.323 /4178 (30 )
=2.5769910/
Now to cool the heat transferred through the combustion chamber material along the passage
which we have formed we need only 7 LPH for cooling the chamber of outside temperature 700C. Now
this is due to the poor conductivity of the material chosen and will melt the inner surface of the
combustion chamber. So this kind of design is a poor design, so to rectify this we can increase the mass
flow of water to a maximum level so as to maintain the chamber wall temperature lower. But this will
divide the combustion chamber wall into two regions, the hotter inner region and the cooled outer region
which may result in the deformation of the combustion chamber. Anyhow for amateur rocket designers
this will not affect the performance in a big way and the design can be implemented.
4.5 DESIGN SECTION SKETCH OF THE HYPOTHETICAL ASSEMBLY
Choose a suitable ignition system for the accounted rocket engine and hence show a sketchof the measurement systems and hypothetical assembly.
Although ignition must be produced when the rich mixture comes in contact with a spark, it is
safer for any amateur rocket builder to choose continuous flame as a source of ignition. Portable LPG
cylinders were available and a small LPG welding torch can be fitted into the combustion chamber by
welding. If an amateur feels that this will be difficult, then pyrotechnic igniters can be used. The
disadvantage is that we need to replace the pyrotechnic igniter whenever we restart the engine. For
pyrotechnic igniter one can simply buy a slow burning fire cracker and simply fix it from the bottom in
such a way that the flame should be in the combustion chamber.
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Performance measuring systems include temperature transducers, pressure transducers and the
thrust transducer. Temperature has to be measured at three places namely at the combustion chamber,
throat and at the nozzle. Temperature can be measured in many ways; out of those ways using a
thermocouple is the best and convenient method for an amateur rocket designer. The thermocouple that
we have chosen is 4.5 inches (114 mm) long and the outer ceramic sealing is 13 mm in diameter.
Provisions for inserting a steel tube of dia 15mm has been made in the combustion chamber. The holes
were not opened so as to avoid the pressure loss. A blind hole of depth 8mm for measuring the chamber
temperature and a hole for measuring the temperature at the throat can be made on the 10 mm thickness
combustion chamber as shown in the figure.
Fig. 4.5-1 Thermocouple provisions in the combustion chamber (All dimension are in mm)
Fig. 4.5-2 Cr/Al Thermocouple with indicator
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The next performance parameter which is to be measured is pressure. This requires a hole in the
combustion chamber, and through the hole, the pressure has to be tapped up and with a suitable pressure
gauge the pressure can be measured. It should be noted that the temperature of the combusted gases will
be high. Most of the pressure gauges will not withstand such temperatures. So a suitable high temperature
pressure gauge has to be chosen for this purpose, Else the gases has to be cooled down. Since most of the
high temperature pressure gauges and digital pressure transducers are of high cost and involve more
complexity in the assembly and calibration we had not made any attempt in measuring the pressure,
though the concept can be understood from the following figure.
Fig. 4.5-3 Pressure tapping provisions
The most important measuring device is the thrust indicator. Thrust is the output given by the
rocket engine. Since this is a force this can be measured by a force sensor. The commonly available force
sensor is the load cell. Nowadays most of the truck weigh bridges use this kind of load cells. A small
range load cell used for that purpose will help us in figuring out the thrust. The following figure shows
the load cell and the assembly can be seen in fig. 4.5-5.
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Fig. 4.5-4 Load cell with indicator
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BLOCK DIAGRAM OF ENGINE ASSEMBLY
Fig. 4.5-5 Block diagram of hypothetical engine assembly
So, with this the design of a Liquid Propellant Rocket Engine with bell contour, choosing
Gasoline and Gaseous oxygen as propellants to produce an overall thrust of 500Kgf is over and the
subsequent sections will discuss about the manufacturing, assembly and testing.
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Chapter V MANUFACTURING AND ASSEMBLY
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INTRODUCTIONThis chapter explains about the entire manufacturing processes which are involved in giving
shape to the designed engine. Since the material we have chosen is EN8 which is comparatively a good
grade mild when compared to any other mild steel grades. Generally many kinds of thoughts may arise
for manufacturing the engine. Out of those the most feasible are the casting and machining.
Casting is a process of producing solid objects by pouring molten material into a shaped mold and
allowing it to cool. Casting is used to shape such materials as glass and plastics, as well as metals and
alloys. The traditional method of casting metal is sand casting. Using a model of the object to be
produced, a hollow mold is made in a damp sand and clay mix. Molten metal is then poured into the
mold, taking its shape when it cools and solidifies. The sand mold is broken to release the casting.Permanent metal molds called dies are also used for casting, in particular, small items in mass-production
processes where molten metal is injected under pressure into cooled dies. Nowadays the mold is also
made in wood and seafoam (Thermocoal). Die casting is a similar process in which the mould is made up
of steel and the molten metal is pressure fed into the mould. This kind of process gives good surface
finish than sand casting and precision parts can also be made. In this method the material required is
exactly the same as in the designed shape. There will be no loss of material when we manufacture in this
process.
Although there will be huge material loss, machining will be the convenient process for this kind
of rocket engine manufacturing. Since it involves complex parabolic contour in the nozzle the shape of
the mould cannot be easily produced. But by computer aided manufacturing (CAM) the required contourcan be more precisely produced. The whole of combustion chamber and nozzle can be easily
manufactured by turning operation itself.
5.1 A BRIEF DESCRIPTION OF THE MACHINE TOOLS
LATHE is the oldest and most common type of turning machine that holds and rotates metal or wood
while a cutting tool shapes the material. The tool may be moved parallel to or across the direction of
rotation to form parts that have a cylindrical or conical shape or to cut threads. With special attachments,
a lathe may also be used to produce flat surfaces, as a milling machine does, or it may drill or bore holesin the work piece.
MILLING MACHINE, In a milling machin