sriram project

8/8/2019 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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Design, Development & Testing of Liquid propellant Rocket engine 2009

i

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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Design, Development & Testing of Liquid Propellant Rocket engine 2009

1

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

sriram project

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