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LIQUID PULSED PLASMA THRUSTER

Spacecraft propulsion system is mainly classified into chemical propulsion system and electric propulsion system. Electric propulsion system is mostly suitable for micro spacecrafts. They always function in vacuum condition and cerate low thrust with high level of specific impulse. Electric propulsion system is mainly divided into three parts, which are Electro thermal propulsion, Electrostatic propulsion and electromagnetic propulsion. The pulsed plasma thruster (PPT) is an electromagnetic propulsion system and is mainly divided in to three sub systems such as solid PPT, Gas FED PPT and liquid PPT. This report is mainly covering background theory about PPT. The aim of this project is to design a liquid PPT engine for micro spacecraft. The report mainly contains the function of liquid PPT and previous research on this topic and performance achieved on previous designs. The design part in this project is about designing a suitable propellant injector for micro spacecraft. The design part contains some suitable designs for injector and the best design is chosen and its performance is calculated.

Contents Page Nomenclature List of figures Introduction Chapter 1 Rocket Propulsion Theory 1.1.1 Introduction to space craft propulsion 1.1.2 Classification of rocket system 1.2.1 The rocket propulsion theory 1.2.2 The magnitude of the rocket thrust 1.2.3 Basic Equation for thrust 1.2.4 The specific impulse of rocket engine 1.3.1 Rocket propulsion system 1.3.2 Chemical propulsion system 1.3.3 Solid rocket engine 1.3.4 Liquid propellant system Chapter 2 Electric Propulsion System 2.1  Electric propulsion system 2.2.1 Electro thermal propulsion 2.2.2 The Resistojets 2.2.3 Arc jets 2.3.1 Electro static propulsion 2.3.2 Electron bombardment Thruster 2.3.3 Filed emission Thruster 2.4.1 Electro magnetic propulsion 2.4.2 Fleming’s left hand rule 2.4.3 Lorenz law 2.5.1 Magneto plasma Thruster 2.5.2 Hall effect Thruster Chapter 3 Pulsed Plasma Thruster 3.1.1 solid pulsed plasma Thruster 3.1.2 the acceleration process in solid pulsed plasma thruster 3.2.1 Gas Fed pulsed plasma Thruster 3.2.2 Mass bit of gas Fed PPT 3.2.3 Measuring the performance of Gas Fed PPT 3.3.1 Liquid pulsed plasma Thruster 3.3.2 Comparing the liquid PPT with gas Fed PPT 3.3.3 The function of liquid PPT 3.3.4 Previous design history of PPT 3.3.5 University of Tokyo’s designed Thruster 3.3.6 Mass bit of injector Chapter 4 Designing a Liquid Pulsed Plasma Thruster 4.1 Designing a liquid pulsed PPT 4.2 Electro spray liquid injector 4.3 Pizo electric capillary injector 4.4 Electro thermal injector 4.5 Mechanical piston injector 4.6 Gas compression Thruster 4.7.1 Choosing the best design out of all suggested design concepts 4.7.2 Designing a mechanical piston injector 4.7.3 Design Calculation Conclusion Reference Nomenclature •		M	Instantaneous mass of the rocket •		U	Velocity •		F	Thrust / net force •		Mp	Mass of propellant •		Veq	Equivalent engine rocket velocity •		Me	Empty mass •		V	Change in velocity •		R	The mass ratio initial mass to mass at burn out. •		 Total mass flow rate of propellant •		 Gravity constant at sea level •		 Total effective propellant mass •		V	Average discharge velocity •		Is	The specific impulse •		Q	Charge of an electron •		EF	Electric field strength •		B	Magnetic field strength •		Vins	Instantaneous velocity of charge particle •		Mbit	Mass bit •		Ue	Exhaust velocity •		C	Measured capacitance •		V0	Initial voltage on capacitor •		E	Capacitor stored energy •			Thrust efficiency •			Density •		A	Area •		P	Change in pressure •		Q	Volume Flow Rate •			Viscosity •		 Length of •		R	Radius •		S	Distance

List of Figures

		Figure 01	Propulsion system Classification 		Figure 02	Chemical rocket thrust equation 		Figure 03 	The solid Rocket Engine 		Figure 04	Liquid propellant Engine 		Figure 05	Ejected exhaust gas from Electro thermal thruster 		Figure 06 	The Resistojet 		Figure 07	Arc Jet Thruster 		Figure 08 	Electron acceleration in filed emission Thruster 		Figure 09	Field emission Thruster 		Figure 10 	Fleming’s left hand law 		Figure 11	Acceleration of ions in electric field 		Figure 12	Magneto plasma Thruster 		Figure 13	Hall effect Thruster 		Figure 14	Acceleration in solid PPT 		Figure 15	Schematic diagram of solid PPT 		Figure 16	Schematic of Gas FED PPT 		Figure 17	Water injector in liquid PPT 		Figure 18	University of Tokyo’s silicon rubber valve injector 		Figure 19	Schematic of silicon rubber injector 		Figure 20	Electro spray 		Figure 21	Pizo electric water injector 		Figure 22 	Electrical heating water injector 		Figure 23	Mechanical Piston water injector 		Figure 24	Gas compression water injector 		Figure 25	Designed mechanical piston injector 		Figure 26	Pressure points at injector

Introduction

In this modern world, we have seen many technological advancements, incredible inventions and ultimate explorations. Many scientists are interested in explorations but very few of them are interested in space exploration. This is because of many complicated facts and technological limitations. To overcome this limitation many space exploration organisations such as NASA and EADS invest huge funds on researches. Apart form these companies many universities around the world establish researches to their maximum potential. This project is also one such university research to refine the space travel. During these space explorations and space travel it is necessary to move the satellite, space shuttle and astronomer to various places to complete successful missions. To move these space vehicles we require different types of propulsion systems and propellants. During the advancement in space exploration, the world witnesses new developments and various new types of propellants and propulsion systems.

In the 13th century Chinese invented solid propellant using the gun powder for fireworks which blasted and propelled the fire work rocket and left the rocket in projectile motion. Until 20th century various solid propellants were invented and used. In 1903 the idea of liquid fuel propellant was introduced by K.E. Tsioikovsky. But in March 16th 1926 an American professor R.Goddard launched his first rocket. Duration of its flight was 2.5s. Even though it flew for only a few seconds, it created a remarkable turning point in the history of space travel. [01]

Due to the disadvantages of Low Mass thrust ratio in this type of propulsion made the scientists to think about another alternative propulsion system, which led to the discovery of a modern technology called electric propulsion system. This system has low thrust but its mass thrust ratio is very high therefore it’s suitable for only space missions. The electric propulsion system is mainly classified into three categories such as Electro static, Electro thermal and electromagnetic propulsion. This project mainly covers the electric propulsion system and its classifications.

The aim of this project is designing a liquid pulsed plasma thruster engine for micro spacecrafts for direction and altitude control. Liquid pulsed plasma thrusters are an electromagnetic propulsion system and ideally come from pulsed plasma thrusters. These pulsed plasma thrusters can be classified into three categories such as solid PPT, Gas FED PPT and Liquid PPT. All these thrusters follow the same principal to accelerate the plasma but they are using different types of propellants in different physical stages. This is done to expand the performance of the thustser and increase the lifetime of thrusters. The solid PPT and Gas Fed PPT have already been designed and their performances have been tested. But liquid pulsed plasma thruster is a brand new technology and still under research by university of Tokyo, Japan. In this liquid PPT development project, my task is to design a propellant injector for the thruster’s ionisation chamber.

To design the best injector for liquid PPT, I have suggested some concepts through my background reading and have chosen the best design from how they have satisfied my design requirements. The performance of the best design is calculated and function of the design is clearly explained in this project.

Chapter 1 I 1.1	.1 Introduction to spacecraft propulsion system

	Propulsion System The main system in a spacecraft is a propulsion system. Which can be classified in three ways, which are 1)	Photon 2)	Rocket 3)	Solar sail

The rocket propulsion system is the ideal system in a small spacecraft. They can be divided in to three main systems •	Thermal •	Electric •	Nuclear

1.1.2 Classification of Rocket System

Rocket

Thermal                                        Electric                                        Nuclear (Under Research)

Electro Thermal        Electro Magnetic       Electrostatic

Chemical   Nuclear     Solar

Plasma Solid      Liquid

Resistojet        Arc jet                                   ion                  collid

Figure 1(by s.vijiprasath) (Propulsion system Classification)[02]

1.2.1 The Rocket Propulsion Theory

A rocket is a device, which carries out matter and energy. The matter is transforms into hot gaseous state then ejects it at a controlled rate through the nozzle at certain direction with high velocity. The main task of the rocket engine is converting the thermal energy to thrust. In this topic we are going to discuss, how the magnitude of a rocket is determined from the velocity and mass rate of flow of the expelling fluid.[03]

1.2.2The Magnitude of the Rocket Thrust

When describing the calculation of the thrust, which is created by hot exhaust gas ejection,it is more complicated to calculate only in atoms. This is due of the addition of pressure forces to the momentum. We can develop the concept of thrust with the momentum theorem of the fluids. Which states that the vector net time rate of change momentum or flux of momentum, through a fixed surface enclosing a steadily flowing fluid equals the vector area integral of pressure or net force acting on the surface.[03]

Thrust force acting on the rocket shell is equal to the sum of all the pressure forces acting on its inner and outer surfaces.[03]

1.2.3 Basic Equation for Thrust

Figure 2 Chemical rocket thrusrt eqation (Source NASA)

The basic rocket propulsion equations are based on Newton’s Law of Motion. Which states that [16] Equation (1) Where M: Instantaneous mass of the rocket U: Velocity of the rocket F: Thrust / net force

Conservation of momentum states that Equation (2) Where Mp: mass of propellant Veq : equivalent engine rocket velocity

Assume that U = 0 We know that mass of the rocket change respect to the time therefore we can write the equation as shown in blow

Where Me : empty mass

Apply both equation (A) and equation (B) as shown below

So the Basic Rocket Equation is given by

This formula can be written in a form called basic rocket equation. Which is TISIOLKOVSKY EQUATION V= VelnR [02] Where V= Rocket Velocity moment. R = the mass ratio initial mass to mass at burn out.

1.2.4 The Specific Impulse of the Rocket Engine

The specific impulse is a most important parameter in rocket fuction. This is use to mesure the efficiecy/ performance of rocket engine. The specific impulse defines by total impulse per unit weight. [07] If the rocket engine has higher the specific impulse that mean, the engine consumed less amount of mass create the thrust and less specific impulse mean that engine has consumed large amount of propellant to generate thrust.

Defining the specific impulse of a rocket engine The total impulse is the thrust force, which changes respect to time. This can be written as in form of                                               Equation (3) Where F: Thrust force But we know that specific impulse is a impulse per unit weight, which can be written as                                             Equation (4) Where : Total mass flow rate of propellant : Gravity constant at sea level : Total effective propellant mass Apply equation (3) on equation (4)

The above equation is the specific impulse equation for rocket engines. [07]

1.3 Rocket Propulsion System

1.3.1 Chemical propulsion The rocket motor is a basic version of the heat engine, and large amount of propellant is burned rapidly under high pressure and the exhaust gas sent through a specially designed nozzle. In this method of propulsion techniques propellant used chemicals. This has an ability to produce high level of thrust on burning. They contains The thrust chamberThe propellant feed system and propellant storage tanks as main components.This propulsion system can be divided in to two main sub systems according to their propellant element physical state. [03]

1.3.2 SOLID PROPULSION A Propulsion System Using Solid Element as a Fuel

The solid Rocket Engine (Figure 3) Sourse (web2.uwindsor.ca/solidrocket.jpg)

It’s generally believed that the Chinese introduced solid propulsion system to the world from the invention of fireworks. The great advantage about this system is very easy to manufacture and simply power the system by using very basic ignition system. For many applications solid propellants giving excellent performance in total impulse to weight ratio.

The solid propellant has an oxidiser and reducer as a main part. In solid propellant, the oxidiser and reducers are mixed together. Then they are mixer sent into the combustion chamber under controlled condition for burning. In order to increase the rate of combustion and fuel efficiency, reducers are in use. Mainly the ammonium per chlorate is used, as an oxidiser and aluminium powder is used as fuel in current solid rocket engines. To increase the rate of combustion iron peroxide function as a catalyst.[03]

1.3.3 Liquid Propellant system A System Using a Liquid as a Propellant Element

Liquid propellant Engine (Figure 4) Sourse (history.nasa.gov/SP-4404/p281.jpg) The main aim of the liquid engine is to transport a payload form one point to another point in space. From the aim we can develop the components as shown in figure 4. 1)	Pay load 2)	Guidance Equipment 3)	An airframe 4)	Propellant tanks 5)	A rocket engine

The main elements of this engine are propellant tanks, propellant feed mechanism, the thrust chamber. The density and weight of the propellant element decide the design of propellant tank. The required thrust in this system is proportional to the weight rate of the flow of propellants.[05] The pressure in the tanks must be constant in order to operate the engine at a constant combustion chamber pressure. These propellants may use either bi-propellants or monopropellants. In bi propellant systems gasoline or alcohol is used as fuel and liquid oxygen or nitric acid is used as oxidiser. Monopropellant system’s liquid has both oxidiser and fuel. This system is stable at normal pressure and temperature. This system simplifies the liquid propellant system’s pumping mechanism. Because of its liquid has both oxidiser and fuel in one single tank.[03]

The Function of the System In liquid propulsion system oxidiser and fuel is stored in two differnt tanks. They are pumped in to the combustion chamber and ignited under high pressure. This pressurised gas mixture will be directed towards the nozzle to provide the thrust to the rocket.The liquid propulsion system contains two major function units[03] a)	the feed system b)	flow control unit c)	thrust chamber unit

The feed system has propellant storage tank pressuring unit. At present most rockets contain two storage tanks so it is essential to have a feed system to pressurise the liquid in the tanks for pumping purposes. The compressed gases such as air or nitrogen is used to pressurise the storage tanks. The main function of the flow control unit is to control the flow rate of oxidiser and fuel mixing in to the combustion chamber. This is because of the rocket operation is for short duration in order. The thrust chamber unit helps propellant to induce a high velocity to atomise, heat, vaporise, and reaction of the propellants before exhaust gases are expanded through the nozzle .[05]

Chapter 2.0

2.1 Electric Propulsion System In the modern world electric propulsion is one of the greatest inventions in the history of aerospace science. The function of electric propulsion is mainly based on using electric power to accelerate the propellant by electrical heating, electric body forces, or magnetic body forces. The electric propulsion always gives low acceleration, they are not suitable under high gravity field and they give excellent performance at the low gravity field. It is due to this reason why all missions with electric propulsion will start from the space, therefore all the operation regarding electric propulsion must be done at vacuum. The electric propulsion system can be classified in to three main sub systems.[02] A)	Electro Thermal Propulsion A propellant is heated by using electrical energy and is thermodynamically expanded and passed as hot exhaust gas through the nozzle to create the thrust.

B)	Electro Static Propulsion Using electrostatic field on charged particle to accelerate and create the thrust.

C)	Electro Magnetic Propulsion Acceleration is created by the interaction of electric and magnetic fields on highly ionised magnetic field on highly ionised plasma. [03] Thrust values of Electric propulsion Type	Specific Impulse 	Thrust / Weight 	Thrust Duration Electro Thermal	300-1500	<103	Years (intermittent) Months ( steady) Electro static 	1000-10,000	<10-4-10-6	Months – Years steady Electro Magnetic	2000-100,000+	<10-4	Years (intermittent) Months ( steady) Ref - Peter J turchi. Chapter 9 Electric Rocket Propulsion System Application of Electric Propulsion in Aerospace Science 1)	Increase speed of satellite to increase the orbit length such as changing the satellite path, lower orbit to higher orbit or geo stationary earth orbit. 2)	A future potential mission is increasing the speed of deep space probes while overcoming the gravitational attraction of the sun and other planets. 3)	Over coming transitional and rotational perturbation in satellites orbits such as north-south station keeping of earth satellite in synchronous orbit or aligning telescope or antennas or drag compensation of satellites in medium-high earth orbits.[09]

The main disadvantage about this system is it always creates low thrust because it takes long time of flight to change it’s orbit. The thrust gives small acceleration and very small increase in velocity for each orbit.

The kinetic power of jet P per unit thrust F can be expressed by flowing equation. P/F = (½ mv2 )/mv = ½ v = ½ Isgo Where M = Mass flow rate V = Average discharge velocity Is = The specific impulse

The ideal power input to the jet is proportional to the exhaust velocity or the specific impulse. The electric propulsion with higher values of Is need more power per unit thrust and a bigger, heavier power supply.[09] The efficiency of the electric propulsion system can be defined from the ratio of the thrust producing kinetic energy of axial component rate of the exhaust to the total electrical power supplied by the thruster, which is

 = Power of the jet / Electrical power input =  Pjet  / IV

From the thrust and specific impulse equation = ½ mv2 / Pe = FIsgo/2Pe = FIsgo/2IV

Where Pe = I x V [09]

2.2.1 Electro Thermal Thrusters (A system using electric energy to heat the propellant element)

Ejected exhaust gas from electro thermal thruster (Figure 5) Source (www.engin.umich.edu/.../images/T-140_firing.jpg)

The main aim of this system is to use electric energy to heat the propellant element and pass through the supersonic nozzle to create the thrust. This system can be classified in to two main sub systems. A)	Resistojet (heating the high resistance Metal) B)	The arc jet ( heating the gas flow by the electric arc discharge)

2.2.2 The Resistojets [06] This device is a basic form of electric propulsion. The propellant element is in gas form and sent through an electrically heated hot surface such as coils of heated wire, heated hollow tubes, over heated blades or over a heated cylinder. For this engine power requirement is 1 w and several kilowatts and thrust can be steady state or intermitted. The elements such as H2,O2, H2O,CO2,NH3,CH4,N2 can be used in this propulsion system. The hydrogen gives high specific impulse for this type of engine because of its lower molecular weight. The disadvantage of using hydrogen in this system is low density, which causes the high pressure storage. There is an advantage of using a liquid hydrazine, which reduces the electric power required to a system by the catalytic decomposition preheats the mixed gases to about 760’c prior to being heated electrically to an even higher temperature. [06] This thruster has 0.01 to 0.025N thrust and gives specific impulse between 280 and 330sec with demonstrates pulsing up to 500,000 times [06]. The Resistojet (Figure 6) Source (www.lr.tudelft.nl/.../img/pact1.jpg ) The Resistojets are suitable for long duration space missions, because of the advantages of using space station waste products such as co2 and h2o as propellants. Resistojets require propellant feed system to supply either gas from high-pressure storage or liquid under zero gravity conditions and gas flow in the system is laminar and mostly heat transfer takes place by the conduction method. Electro thermal propulsion system commonly experiences intermittent heat transfer from the heating element to the propellant by conduction and radiation losses from the thrusters. The maximum achievable temperature of system is 2700k which is controlled by the highly restive materials such as rhenium alloys, platinum, molybdenum alloys or tungsten alloys. The boron nitride is more effective for high temperature electric insulators.

Thrust efficiency of resistojets between 65-85%, is depending up on the following factor a)	The propellant b)	The exhaust gas temperature c)	Numerous design details

The specific impulse of the resistojets is decided by the properties and character of the propellant used in the system, the maximum allowable temperature in the chamber and nozzle surface. [07]

2.2.3 Arc jet Thruster Function of the system When an electric power is supplied to the system, the cathode tip is super heated up to 2000 to 2300 by the arc interaction, which makes pre-heats the propellant gas. The inner walls of the nozzles are also super heated to 10,000K-20000K by the radiation from the arc. The hot gas in the arc mixes with the outer annular and is sents through the super sonic nozzle to create the thrust to the system.[06]

Arc Jet Thruster (Fiqure 7) Ref - www.daviddarling.info/encyclopedia/A/arcjet.html

Arc jet thrusters have higher performance than resistojets, this is because they don’t have limitation in gas temperature and specific impulse like resistojets. Using arc for direct heating avoids these kind of problems. Normally arc stretches between the central cathode and the coaxial nozzle,[06] which is known as the annular anode of the system. Both cathode and anode require electrical insulation and they are capable of withstanding high temperatures. As shown in the above diagram arc must be placed at the entrance of the nozzle. The plane of anode attachment has an ability to move up or down according to the mass flow rate of exhaust and applied voltage. In this system small amount of gas is superheated by the arc and then ionised, the remaining gas is cool and prevents the nozzle from the super heated jet. The increment in velocity and Mach number is only caused by the heat released in the gas upstream of the nozzle. There are some several disadvantages in this system, the created arc is inherently unstable so there is a chance for the formation of pinches and wiggles but they can be stabilised by the external electric field. Power supply to the arc jet system determines the specific impulse created by the thruster and arc jet temperature. [07]   The lifetime of the arc jet depends on the maximum allowable temperature of the heated material.

2.3 Electrostatic Thrusters Electro static propulsion accelerates an ionised propellant gas by the application of electric field. The stability depends on the electrostatic field strength. They can only operate in the vacuum condition. They require all the charged particles to have same sign, in order to the move the charge particle in the same direction. In Electrostatic propulsion electrons function as a propellant element and they are very easy to produce and easy to accelerate through the high velocity but they are extremely lighter particles so even if their velocity is high the momentum is negligible. To avoid this, electrostatic thrusters use charged heavy atoms such as positive ions, charged colloids, usually small liquid droplets which in turn are at least 10,000 times heavy as ions.[07]

The Electrostatic Thruster an be Divided in to Three Main Sub Systems 1)	Electron bombardment thrusters Gaseous propellant such as xenon or mercury produces positive ions by bombarding.

2)	Ion contact thrusters Passing the propellant vapour such as cesium, through a hot porous tungsten contact ioniser to produce the positive ions.( still under research)

3)	Filed emission or colloid thrusters Passing droplets through intense electric field charges (tiny liquid droplets of propellants either positively or negatively charged)[07]

2.3.1 Electron Bombardment Thruster

Electron acceleration in filed emission Thruster (Figure 8) (Source  www.daviddarling.info/images/electron_bombard)

In the electron bombardment thruster, the filament inside the chamber functions as an electron producing device. From this device electrons are discharged at low chamber pressure. In this system electrons are attracted towards the cylindrical anode and they are controlled by the a magnetic field which is around the system. The ionisation process on this system relies on the electron energy, electron density and density of propellant atoms in the ionisation chamber. The chamber contain of positive ions, electrons, and neutral atoms. The strong electric field in the vicinity of the pair of grid electrodes at the exit of the chamber extract and accelerate the positive ions through the openings in the grids.[07] The potential of the cathode in the chamber is being negatively biased on both the inner grid electrode and the opposite wall of the chamber to prevent the electron loss. In the past cesium and mercury were used as a propellant element, due to some disadvantages such as being hazardous and some times leaving wastage on the space craft which causes reaction with other materiel. They have stopped using cesium as a propellant element in the system. Currently xenon, the inert monatomic gas with high atomic weight is being used as a propellant element in this system. [07]

2.3.2 Field Emission Electric Propulsion Thrusters

This propulsion system is different from the other electrostatic propulsion systems, because of the application of electrostatic field to accelerate the positively charged atoms. In this system they use a liquid form of metal such as ceseum or indium as propellant element. Because these metal have high atomic weight low melting point and low ionisation potential. This system gains thrust by application of a strong electric field to pull liquid propellant off a tungsten needle.

1,000 to -6,000 V potential is required for the accelerating electrodes, generating a field at the tip of about 1 V/nm. The generated thrust depends on whether ions or droplets are emitted from the needle tip. The emitter in this system consists of two metallic plates and a small propellant reservoir.

Field emission Thruster (Figure 9) Ref - http.engin.umich.edu/.../FeepSchematic.jpg

When the two-emitter halves are tightly clamped together, the propellant flows through tiny channel, forming a free surface at the exit of the slit. The radius of curvature is in the order of 1 µm. when a strong electric field is formed by giving a voltage difference between the emitter and an accelerator electrode located directly in front of it. When a free surface of the liquid metal approaches a condition of local instability due to the combined effects of the electrostatic force and the surface tension. it foams the "Taylor cones". When the electric field approaches a certain value, the atoms at the tip spontaneously ionise and a thrust-producing ion jet is extracted by the electric field, while the electrons are rejected in the bulk of the liquid.[07] For the needle design (1d), the emitter is a single needle with a pool of indium melted in a reservoir with an external heater. When we compare colloid Thruster with an ion thruster, colloid thruster has some advantages such as requiring lower power to thrust ratio, higher thrust efficiency and higher thrust per unit area. And also there are several disadvantages on this system like electrical leakage from needle to accelerating electrode and colloid particles; they have non-uniform size and some problem with their charge intensity.[07]

2.4 Electro Magnetic Propulsion

The basic concept of a electromagnetic propulsion system is propellant element changes to a plasma state and is accelerating it by applying both electric and magnetic fields. First of all I would like to explain about plasma, ions, electrons and ionisation process which are very essential parameters to learn. Plasma is an electrically neutral gas and contains equal number of positive and negative charges, which is the fourth state of matter and they have an ability to conduct electricity at high temperature levels such as 5600K. We consider the ion simply as an atom or molecule which carries out electric charges. The ionisation process will take place in atom or molecule by adding or removing an electron. The sign of the ion depends on whether they gain or lose an electron from the atom. If they lose one or more electrons they are positively charged ion and if they gain one or more electron it is known as negatively charged atom.

According to the electromagnetic theory, whenever a conductor placed in a magnetic field carries current, a force will be exerted on the system, this is due to Fleming’s left hand rule.[07],[15] 2.4.1 Fleming's left hand rule

Fleming’s left hand law (Figure 10) Ref http://upload.wikimedia.org/wikipedia/en/c/cb/Left_hand_rule.png

When a current carrying conductor is moved in a magnetic field the conductor reacts with the magnetic field causing the conductor to move outwards. The direction of movement can be predicted as shown below. Thumb – Thrust force First finger – Magnetic field direction Second finger - Direction of current (positive to negative)[16]

Electro magnetic propulsion system prevent charge accumulation because they accelerate plasma to gain a thrust, the plasma consist of both positive and negative charge in equal amount. Another advantage in this system is high thrust density, which is normally 10 to 100 times greater than electrostatic ion thruster. The electromagnetic thruster is classified in to many ways such as Electro dynamic, Magneto plasma dynamic (MPD), Hydro magnetic, Magneto plasma, Pulsed-plasma, Pinch engine, Hall accelerator, Lorenz force accelerator, Pulsed co-axial thruster, etc. [16], [06]

Acceleration of ions in electric field (Figure 11) Ref. –[15]

When a highly electrically conductive fluid such as ionised gas is subjected to electric field E and magnetic field B, and if both fields are perpendicular to each other to the moving fluid with velocity of u as shown in above diagram. Here current density j driven by the electric field interacts with B to provide a steam wise body force F = j x B that causes the acceleration to fluid along the channel. The above situation can be simply explained with the Lorenz law [15].

2.4.2 Lorenz Law

The Lorenz law states that when a charged particle moves in both electric field and magnetic filed it will experience a force. When a charged particle is in an electric field region it will experience a force, which is given by [17] FE=   q x E  Where Q = Charge of an electron E = Electric field strength

A force due to magnetic filed is given by: FB=  q x v  x B Where B = Magnetic field strength V = Instantaneous velocity of charge particle The addition of both equations gives the standard equation for the Lorenz force (F). Which is        F = q x (E + v x B)

Electromagnetic propulsion mainly creates plasma by applying a high current discharge through the propellant. [17]

Ionisation Process The ionisation process takes place when an electrically charged atom or molecule looses or gains electrons. [10]. When an atom or molecule loses one or more electrons, they are positively charged and when they gain one or more electrons they are negatively charged. [07]

Thrust Mode	Steady State	Pulsed Magnetic field source	External coils or permanent magnets 	Self induced	Self – induced Electric current source	Direct – current supply	Capacitor bank and fast switches Working Fluid	Pure gas, gas mixture, seeded gas, or vaporised liquid. Pure gas of vaporised liquid or stored as solid Geometry of path of working fluid	Co-axial rectangular, constant or variable cross section 	Pinches by field, ablating plug co-axial, other Special Features	Using hall effect for changed electric field Information from –Rocket propulsion elements, George P. Sutton

2.5 Types of Electromagnetic Propulsion System

(A)	Magneto plasma dynamic thrusters (B)	Hall Thruster (C)	Pulsed plasma thruster

2.5.1 Magneto plasma dynamic Thruster Magneto plasma Thruster (Figure 12) (Ref. – Princeton University journal)

The above diagram shows the magneto plasma Thruster, which is characterised by the coaxial geometry constituted by a central cathode and annular anode. The gaseous propellant is ejected to the upstream portion of the channel, where the gas is ionised by applying a uniform electric arc in the inter electrode gap. The arc current used in this system is very high and induces a magnetic field, which is enough to exert the desired axial body force on the propellant flow. They directly accelerate the plasma towards the cathode tip. This expansion and acceleration of the plasma causes the thrust to the system. These thrusters can provide specific impulse of 1500-8000 sec with the thruster efficiency of 40%. They reach high efficiency when using high power levels. The steady state version of the MPDT is regarded as option for high power propulsion. The MPDT is especially suitable for energetic space missions. Because they require high thrust level for the carrying cargo to other planets. Another version of MPDT is lithium Lorenz force accelerator which has multi-channel hollow cathode and use lithium as propellant element to reduce the cathode erosion which helps to increase the thruster efficiency at moderately high power level. This accelerator has 500-hour erosion free operation and can provide a thrust of 12.5N, specific impulse of 4000s. The thruster efficiency of this system is around 48%. So far no other electric thruster is operated with this power level. The special feature of MPDT is, they can be operated in quasi-steady pulsed mode by using flat top high current pulses. [15]

2.5.2 Hall Effect Thruster

Hall Effect is when electromagnetic accelerator is operated with low plasma density or high magnetic field, the current passes through, it will divert from strict alignment with the applied electric field to acquire a component in the E x B direction. This system allows the positive ions to accelerate freely down stream by applying electric field. This thruster can achieve efficiency up to 50%. Hall effect Thruster (figure 13) (Ref. – Princeton university journal) The above diagram shows the schematic cross-section of hall thruster. In the system the propellant is xenon gas which has high molecular mass and requires less energy for the ionisation process. In this case xenon gas is ejected through small holes in the anode. The ejected propellant is ionised by collision with high-energy electrons, these high-energy electron normally have 10–20 eV amount of energy. Then the created ions are accelerated by the electric field, which is created, between cathode and anode. They create ring type electron clouds at the exit the by method of electrons tapping the magnetic field and ions are accelerated towards it. When ions are leaving out of the thruster they pull equal number of electrons with them and creating plume with no net charge. Main function of magnetic field in this thruster is ensuring that the discharge power is only applied to accelerate the xenon gas ions and not the electrons. They require being strong enough to deflect the Low Mass electron and not the heavy xenon ion. The mass flow rate of this thruster is around 5mg/s and voltage of 200 to 300v yielding the plasma to the exhaust velocity of 16,000ms-1and thrust range is around 40-80mN, at efficiencies of about 50%.[15]

Chapter 3 Pulsed Plasma Thruster

Pulsed plasma thruster is an electromagnetic propulsion device. They use electrical energy to charged the capacitor and create strong electric arc and discharge across the propellant for ionisation process. The discharge ablates the propellant and converts them in to Plasma State. The created plasma accelerated by the high electric field and magnetic field escapes with the high exhaust velocity to create the thrust. The electromagnetic force used to accelerate the plasma is known as Lorenz force.

Types of Pulsed Plasma Thrusters (1) Solid Pulsed Plasma Thruster (2) Gas FED Plasma Thruster (3) Liquid Plasma Thruster

3.1.1 Solid Pulsed Plasma Thrusters

The ideal version of pulsed plasma thruster is solid PPT, which was first introduced in 1950s by a group of United States, Europe and Soviet Union. In 1964 the first PPT was tested in Soviet Zond 2. The solid pulsed plasma thruster employs Teflon as the propellant element. Because of its thermal stability and it does not react with other elements easily. The Solid PPT is a completely self-contained propulsion system and does not require any other supporting device, which has only one moving part as Negator spring. The whole system has its own thruster, propellant, propellant feed unit and power processing unit. The most important advantage of PPT is it gives high specific impulse at lower bus power level, the solid PPT only requires 28V as a power input. The advantage of using solid propellant is it makes the system very simple and it does not require propellant storage tank, propellant feed system and vales. The thrust created by the PPT can be controlled by the firing frequency and their trust level is around 1.5 milli Newton. They have a specific impulse range of 800-1200s. [08]

3.1.2 The Acceleration Process in Solid PPT Acceleration in solid PPT (figure 13) Ref -http://www.busek.com/downloads/MicroPPT.pdf The above diagram explains the acceleration process of plasma, when plasma is formed in between the cathode and anode, they produce the current sheet. When current passes through the conductor, they induce magnetic field around the electrode region. According to Fleming’s left hand rule and Lorenz force the plasma will accelerate to exhaust velocity. Schematic diagram of solid PPT (figure 14) http://www.daviddarling.info/encyclopedia/P/pulsedplasmathruster.html

In the above diagram the thruster consists of a propellant Teflon bar placed in between the electrodes and pressed toward the ionisation unit by the negator spring. The function of the spring in this system is to feed the propellant element to the ionisation chamber as soon as it is consumed. Power processing unit in this thruster charges the capacitor to a voltage from 500v to 2000v by using unregulated power spacecraft bus and also supplying power to spark plug for the ignition process. For each discharge the capacitor only uses 5 to 50J, once discharge is ignited they ablate the propellant from the face of propellant bar and ionise it. The ionised propellant gas is converted in to the Plasma State and send into the ionisation chamber where acceleration of plasma takes place. The formation of Lorenz force causes the plasma to accelerate to the exhaust velocity. The thermal expansion of non-ionised Teflon vapour also generates some lower amount of thrust. The above combined thrust effectively gives the specific impulse of 800 to 1200s.[08]

In the early 1960’s United States started the development for the Teflon PPT. In 1970s LES 6 was tested and provided the thrust of 26N and specific impulse of 312s. Which were been used in station keeping role of satellites for 5 years. The success mentioned above in the mission made PPT being considered for other space missions. The solid pulsed plasma has given good performance for nano satellites. They are simple and has high level of reliability and the specific impulse created by the thruster is well enough for nano satellites. [08]

3.2.1 Gas Fed Pulsed Plasma Thrusters

The gas fed pulsed plasma thrusters is another version of electromagnetic propulsion. Which has same acceleration unit as solid PPT. but gaseous propellant element is supplied to the ionisation chamber to produce plasma. The disadvantages from solid PPT such as poor performance characteristic, contamination and non uniform ablation causes modification for solid pulsed plasma thruster so scientists decided to remove the solid Teflon and tried to substitute it with gas, which made way for the invention of gas fed thrusters. The gas fed thruster is a new version of electromagnetic propulsion and has a similar acceleration system as solid pulsed plasma thrusters but performance wise gas fed thruster gives an excellent performance when compared to the solid pulsed plasma thruster. They have an ability to provide the specific impulse of 5000s and thruster’s efficiency of 57% (the results from the xenon gas) and only require 130J as discharge energy.[12] Normally argon or xenon gases are used as a propellant element. Normally this thruster operates in unsteady pulsed mode, which is around 3s per pulse. The highest performance for gas fed thruster reported for argon gas. The yield efficiency was 50% and only requires 5J of capacitor energy. An impulse bit was 32Ns and mass bit of 0.2g per shot. Both solid PPT and gas Fed PPT has advantages such as, they always give high specific impulse at very low bus powers, contain very simple form of discharge circuit unit, high instantaneous thrust density and specially they have repeatable impulse bit. But APPT has following advantages over the gad fed PPT, a propellant feed system and pressurising unit are not needed, solid Teflon is more stable, and they don’t require storage system, there is no propellant leak and they don’t have any moving part at all. The special feature of gas fed PPT is: they don’t have space contamination issues, wide range of propellants can be used, they can provide much wider range of achievable specific impulse to the system, higher potential of scalability to low discharge energies and small length scale, repeatable better performance than the APPT due to the latter’s reliance on  surface conditions for the ablated mass production, lower susceptibility to the production of slow neutrals.[11] (Information is from Princeton University, USA)

Recently gas fed PPT with a high repetition rate and high propellant utilisation energy has been developed. They have an operational lifetime up 109 shorts corresponding to a total impulse in excess of 105Ns.

The gas fed thrusters shown in the diagram below is coaxial in geometry and here gas injected at the base of centre electrode which act as the anode of the system and mounting holes around the circumference of outer electrode acts as the cathode of the system.

Schematic of Gas FED PPT (figure 15) Ref- [12]

In above gas fed thruster a steady rate power supply charges the capacitor of the thruster, which discharges the power between two co-axial electrodes where gas is injected before the pulse. When the gas discharge uniform initiation and current sheet form are accelerated due to the Lorenz force created by the current caring plasma, where plasma act as a conductor and self induced magnetic field due to this current. Then gas is accelerated to high exhaust velocity. For this operation, here only the system use 240V of voltage and 10KA of current and required capacitor energy is around 1-10J. The gas fed PPT normally has low propellant utilisation energy this is because of the fact that the fastest available space qualified gas valve with long enough life (>107 pulse) has an open time not shorter than 1ms. As the time to fill up the discharge chamber is only the in the order of 100s, much of the propellant cannot be utilised effectively over the life of the device.[12] But this problem is substituted by using state of the art pulse power technology, which increases the utilisation efficiency up to 90%. The recently designed PT5 (shown in above diagram) is relatively large and contains stainless steal coaxial set of electrodes. They consist of four semi-conductor type plugs and they are mounted on the inside of the outer electrode. Here main discharge is operated by the capacitor, which has a capacitance of 130 or 270. They have an ability to charge to 250v giving a maximum energy per pulse of 4 or 8 J. the coaxial geometry of the system has several advantages; they increase the spark plug life time and the uniform discharge symmetry. To create uniform and symmetric discharge, four or more spark plug is fired simultaneously. When only one spark plug is fired instead of four, it reduces the impulse bit by as much as 40% where measurement was conducted by the NASA JPL.[11]

3.2.2 Mass Bit of Gas FED PPT

Generally argon is the propellant in gas fed PPT. They are supplied to the ionisation chamber via a valve, choking orifice, and plenum. The valve is placed on approximately 10cm from the thruster propellant injection ports. Their function controlled by the control module and the timed by a signal generator. To achieve a steady mass flow rate to ionise the gas, the valve was open for 45ms and the thruster was set fire 45s later the valve opens.

The mass flow rate to the ionisation chamber is controlled by changing the pressure in the plenum upstream of the sonic orifice. The sonic orifice made by the copper plate and has a diameter of only 0.36mm. The mass flow rate can be set in to steady after 20ms of the valve opens. [11]

Measuring the Performance of Gas Fed Pulsed Plasma Thrusters

The performance of Gas fed PPT can be calculated by the produced impulse due to the measured amount of stored energy and mass. The efficiency of the PPT can be given as the ratio of the directed kinetic energy to the initial stored energy.[11]

Which is given by the

η = (1/2mbit ue -2 )/ (1/2CV02) = I2bit / mbit CVo2 [12]

Where mbit	Mass bit ue	Exhaust velocity C	Measured capacitance V0	Initial voltage on capacitor

Note For a PPT, the thrust to power ratio is the same as the impulse bit to energy bit ratio. So

T/P= Ibit / mbit go     =  Ibit / ( 1/2CV02)

The specific impulse is given by the amount of impulse provided by the thrusters per unit weight on earth of the propellant used in the pulse, which is given as Isp = Ibit / mbit go      =  ue/ go

Gas fed thrusters generally experience problem in delivering the propellant to the ionization chamber, the gas valves on the system are not reliable enough to supply the propellant directly from a central storage location to individual thrusters. The gas vales which eventually leak will draw little controversy. To avoid this problem separate plenum is introduced to be used as a short-term gas storage facility during manoeuvres. They are filled with a quantity of propellant deemed to appropriate for given manoeuvres.[11] The plenum uses high cycle life valve to deliver gas propellant to thrusters, this valve may be allowed high leak rate when closing the valve. There is a disadvantage in using the liquid valve because they can be vaporised when it impacts the inner surface of the PPT centre cathode but they have an ability to reduce the leakage of its surface tension.[11]]

The performance data for fixed energy of 5.3+/-0.2J mbit ( g)	Ibit (Ns-1)	Isp (s)	 (%) 0.2	32.6	16,600	49.9 0.2	31.1	15,900	45.5 0.3	28.8	9800	25.9 0.4	28.9	7380	20.0 0.5	29.3	5970	16.1 0.6	30.1	5120	14.2 0.7	29.2	4250	11.5 0.8	28.4	3630	9.6 1.0	29.5	3010	8.4 1.5	29.8	2030	5.6 2.0	29.7	1770	4.2 +/-2%	+/-8 %	+/-8.3%	+/-9.2%        (Information from Princeton University, New Jersey, USA)

3.3.1 The Liquid pulsed plasma Thruster

Introduction to the Liquid Pulsed Propellant System

The previous problems in solid PPT, gas FED PPT such as the thruster weight due to the propellant, contamination by PPT exhaust which is only available in solid PPT where Teflon is a propellant element due to the fact of its stability in vacuum. Teflon is a component made by the carbon and florine, the carbon in the component causes carbonisation to the spark plug and reduces the lifetime and florine causes damage to the solar cells and optical measurement system in spacecraft. The gas fed thruster has good performance but they are more complicated to operate due the complexity in using vales and compression unit and gas quickly escape in the vacuum and cause propellant leakage in space. Above problems in both systems made the way to think about liquid pulsed plasma thruster, where the concept of pulsed plasma thruster remains constant and here liquid is introduced as propellant element. [14]

Performance of Liquid PPT Specific Impulse – up to 4300s Thruster Efficiency – 13% Required capacitor Energy – 20J Unit thrust per mass bit - 1N/s

Using liquid as Propellant The advantage of using liquid as a propellant is, the mass flow rate of liquid can be controlled according to the injector orifice size and late time vaporisation in vacuum when compared to gas, which makes most of the ejected liquid go under the ionisation by the discharge current. The main reason for using water as propellant is it doesn’t contain carbon as atoms as the result of ionisation it doesn’t produce carbon atoms so the carbonisation of the system can be avoided and it does not contain reactive elements such as florine.[14]

3.3.2 Comparing the liquid PPT with Gas FED PPT

Both liquid and gas fed PPT has similar function as the propellant is injected in to the Ionisation chamber. When the gas is injected in to the ionisation chamber within short period of the time it fills the inter electrode region and escapes in 100s. the system require fast acting valves to use the propellant efficiently. However its a bit hard to develop fast active vales with long lifetime and high reliability, normally space missions take place for long time. In order to avoid this problem gas fed PPT with high repetition rate discharge has been proposed and dramatically improves its propellant utilisation efficiency. Gas fed PPT is not suitable for micro thruster because one set of operation requires 20-pulse and capacitor energy of 100J. The function of liquid propellant is different from the gaseous propellant because when the liquid is ejected in to the vacuum it vaporises and escapes from the ionisation chamber more slowly than the gas. Therefore fast acting ejection valve is not required in the liquid PPT. As the result of vaporising mass regarding the temperature and vapour pressure gave 20% of water would be vaporised in 10ms since it is ejected in to the vacuum.[13]

3.3.3 The Function of Liquid Pulsed Plasma Thruster Water injector in liquid PPT (figure 16) Ref - [13]

As shown in above diagram liquid PPT consists main parts as (1)	Bank capacitor (2)	Injector (3)	Igniter (4)	Electrodes

In the system capacitor gets charged form the solar cells of the spacecraft and supply the high current to sparkplug for ionisation. Once the injector eject small amount of water in to the ionisation chamber, the spark plug discharges high current and ionise the little amount of water and convert it into the highly conductive plasma. When plasma conducts current from cathode to anode, they induce magnetic field around the system. Induced magnetic field, electric field causes the plasma to accelerate to the high exhaust velocity from the thrust. The generated thrust is very low which is around 1N/s per mass bit but the specific impulse is around 4300s.[13]

3.3.4 Previous Design History of Liquid Pulsed Plasma Thruster

The recent development in liquid propellant pulsed plasma Thruster has been done by University of Tokyo, Japan. They introduce valve injector to inject the water propellant in to the ionisation chamber. Fist they tried to introduce piezo electric device to inject the water, which has no shut off valves and seals the water by surface tension and pizo electricity is the ability of crystal and ceramic materiel to generate voltage in reponse to applied mechanical stress (wikipedia). Later they found that the piezoelectric capillary injectors could function in atmospheric condition. In vacuum it doesn’t function due to the propellant leakage.[14]

University of Tokyo’s Valve Injector

University of Tokyo’s silicon rubber valve injector (figure 16) Ref [13]

The above diagram shows the University of Tokyo’s design of a valve injector. They eject the liquid form an orifice by producing inner pressure. The (a) is a normal state of the valve. The spring in above diagram presses the closing boss downward to the orifice. Flexible Silicon rubber functions as the sealing equipment. The actuator in the system moves the boss upwards, which makes the valve to open and eject the water by the inner pressure. They found that mechanical vales are not suitable for this thruster due to the lack of simplicity. Their designed valve functioned in vacuum condition and ejected the mass shot of 3g.

The Designed Thruster Schematic of silicon rubber injector (figure 18) Ref - [14]

Above diagram shows designed propellant injector for liquid PPT, where they employ 20J capacitor and parallel plate electrode. In order to increase the inductance per unit length, the parallel plates are in use, they are one of the important facts to decide the thruster efficiency. The electrodes made by carbon are 10mm in height 5mm in thick and 50mm long. The estimated inductance per unit length of the electrodes was 0.9H/m. The electrodes are covered with side walls and a back plate made of silicon. The function of these slide glasses is to prevent the electrodes and injector from the undesirable discharges.

The electrodes are connected to the mica plates capacitor and charged up to 3KV. The bank capacitor of 3 to 4.5F was used.

3.3.6 The Mass Bit of Injector Mass bit ejected by the injector can be calculated by the following equation M = 2E / g2 I2sp        (1) Where M = Mass shot bit G = Gravity constant E = Capacitor stored energy = Thrust efficiency

For their design they achieved, the specific impulse of 1000s, Thruster efficiency of 10% and 40g of mass shot. The mass shot was calculated by dividing the weight difference with number of injection. From their design they estimate the mass flow rate from 2.7- 38g.[14]

Chapter 4

4.1 Designing a Liquid Pulsed Plasma Thruster In this project my task is to design a suitable propellant injector for liquid PPT. In order to design an injector I have come across with some basic design requirements for injector. All these requirements are made from considering the previous design factors. Design requirements 1.	Low leakage 2.	Compactness and lightness 3.	Low power consumption 4.	Supplying a few micro-grams of the propellant 5.	Fast gating 6.	Function In vacuum

I establish my design work by calculating the required mass bit for the injector. Where injected mass bit is an important parameter to measure the performance of the thuruster. In order to find out the mass bit I have used the equation (1).

Folowing values to calculate the mass bit were assumed from the previous design on liquid pulsed plasma thrusters. Where first of all I have chosen the specific impulse of 4000s, in the pervious design on the liquid PPT they have found that 4300s-3900s is a reasonable achievable specific impulse for liquid PPT due to that fact I chose the number in between the range as my supervisor advised me to do so. The previous design has achieved the thrusters efficiency of 13% so I used same the value as they have used .University of Tokyo tested the design by using the capacitor energy of  20J, for my design I used the 15J capacitor, which is a reasonable capacitor energy for liquid PPT.

Assumed values to calculate the mass bit, by considering the previous design data. Specific Impulse            4000s Capacitor Energy          15J Thrusters Efficiency     13% Gravity constant           9.81m/s2

Calculating the mass bit by equation (1) M = = = 2.53283x 10-7 Kg so required  mass bit for the injector is                                                   2.53283x 10-7 Kg To design the injector with above required mass bit following designs were suggested. All these designs concepts were checked in accordance to the design requirements and each concept was analysed.

Suggested Design Concepts for Propellant Injectors

1.	Silicon Rubber Valve 2.	Electro spray 3.	Piezo Electric Capillary injectors 4.	Heating 5.	Mechanical Piston 6.	Gas Pressure Compression

Silicon rubber vale The University of Tokyo has developed this concept. The function of the concept has already been explained in the previous chapter.

4.2 Electro Spray technology Water Injectors It uses the electro spray to inject water into ionisation chamber, where electro spray functions as the injector. The highly electric conductive liquid is pushed by applying an electrostatic charge, where you can get very small amount of mass flow which can be controlled by applying a different voltage. [19]

Electro spray (figure 19) Ref - www.ohsu.edu/proteomics/images/electrospray2.jpg

The above Electro spray uses a every sharp needle with tiny hole, liquid is pumped in to the needle and high current and voltage is supplied to the tip of the needle and forms a chare corn which is known as Taylor cone. At the tip of the cone, liquid becomes fine jet and becomes unstable and breaks in to droplets. The above elector sprays injector were discussed in previous year Liquid PPT project.

4.3 Piezo Electric Capillary Injectors

This type of technology has been introduced in inject printers, where the system uses thermal expansion method to eject the water. Normally pizo electric dice is made up of material such as crystal or ceramic. When we apply a high voltage to the device it generates strain force and pushes the water container. In the space situation using liquid water is impossible due to the lack of gravity the water will be sucked out from the water container due to that reason ice cube are used.[18]

Pizo electric water injector (Figure 20)

The above system consists of electric filament and pizo electric device as main parts. When we supply the current to electric filament, it melts some amount of ice cubes, simultaneously pizo electric device gives some stain to the ice container that cause external pressure in container walls and ejects the water into the ionisation chamber. The ejected amount of water has very low mass flow rate and function of above injector is possible in vacuum condition.

4.4 Electrical Heating Method

This concept is very simple and uses the vacuum condition state to eject the water. The system employs ice cubes as propellant and melts the ice by electrical heating method which allows liquid water to escape by the vacuum condition to the ionisation chamber.

Electrical heating water injector (Figure 21)

The above system only consists of filament coil as a main part. When the filaments is subjected to electrical force it generates heat and melts the small amount of ice cubes and convert it into water. Due to the vacuum condition water escapes towards the electrode where high current ionises the water and converts it into the highly conductive plasma and accelerates to high exhaust velocity to create thrust.

4.5 Mechanical Piston Injector

The mechanical piston injector is very simple version of propellant injector which has a cylindrical shaped water container connected to the piston. At the other end of the water container it has small orifice which is specially designed to inject the small amount of mass flow.

Mechanical Piston water injector (Figure 22) The above diagram shows a mechanical piston water propellant injector for liquid pulsed plasma thruster. The injector is designed to inject the required mass bit of 2.53283x 10-7 Kg. The water containing tank in above system is connected with the small hollow tube with the very little radius, where the tube functions as the orifice. The insulation unit prevents the propellant from freezing and keeps the water in standard temperature. The force generator produces very small amount of force to move the piston in certain velocity to inject the required mass bit of water. When the water is ejected into ionisation chamber where the igniter produces high current to ionise the water propellant and convert it into the plasma stage. The plasma conducts the current from anode to cathode and form strong electric field and magnetic field. The induced electric, magnetic field accelerates the plasma into high exhaust velocity to create the thrust.

4.6 Gas Compression Propellant Injector

This type of propellant injector is already being used in chemical liquid propellant rockets in the propellant feeding unit they supply the oxidiser and fuel in controlled rate to the combustion chamber. In liquid plasma thruster, the gas compression injector uses helium gas to compress the water and injects it into the ionisation unit.

Gas compression water injector (Figure 23) The above diagram shows a designed water injector for liquid pulsed plasma thruster. They are mainly made up with two tanks and a mechanical valve. In the above system one tank is fully filled up with the helium gas and another tank is filled up with water. The insulation unit covers both tanks. They maintain the standard temperature and pressure of both helium gas and water. The mechanical valve controls the volume of helium gas entering the water-containing tank. Once the helium gas is released they compress the water and displace the water in controlled rate volume. There is a small hollow tube connected at the tip of propellant tank, which is directly facing the ionisation unit of the system. The hollow tube releases the small amount of propellant according to the pressure drop due to the gas compression. The ejected propellant is directed towards the ionisation chamber. Where the ionisation of water takes place and accelerates to the exhaust velocity.

4.7.1 Choosing the best design out of all suggested designs

The table below is to find out which is the most important requirement when comparing with other requirements. In the table each requirements are labelled 1 to 7 along the columns and rows. Here each and every requirement compares with the other. When comparing if one requirement is important than the other it will get 1 points if not will get 0 points. Along the row their points are added together.

1	2	3	4	5	6	7	Total 1		1	1	1	1	1	1	6 2	0		1	1	1	1	1	5 3	0	0		0	0	1	0	1 4	0	0	1		1	1	1	4 5	0	0	1	0		1	1	3 6	0	0	1	0	0		0	1 7	0	0	1	0	1	1		3

Where 1.	Operation In Vacuum 2.	Supplying few micrograms of water 3.	Lower power consumption 4.	High reliability 5.	Fast gating 6.	Compactness and Lightness 7.	Life time

In order to get the best design from the above table, each requirement was arranged and the achieved total score multiplied by 10 to get perfect range of scores. Then each of the designs is compared by how they satisfy the design requirement and their percentage is rated as shown below 0	Not satisfied 1	Partially satisfied 2	Fully satisfied

Design score = Rate x requirement weight

Requirement	Weight	D 1	D 2	D 3 Operation In Vacuum Ability to work in vacuum condition	60	120	120	120 Supplying few micrograms of water Ejecting the required mass bit in to the ionisation chamber	50	100	100	50 Lower power consumption Power required for the thruster operation	10	20	20	10 High reliability Must be able to produce mass bit with high rates when required 	40	80	80	40 Fast gating Frequency of mass bit, how many bits can be produce in certain time	30	60	60	30 Compactness and Lightness Thruster should be suitable for micro Space craft	10	20	10	20 Life time Must be suitable for long space mission 30	60	60	60

Where D 1 – Mechanical Piston injector D 2 - Gas Compression Injector D 3 – Thermal Injector

The total achived score Injector	Total Score Mechanical Piston Injector	460 Gas Compression Injector	450 Thermal Injector	330

Why mechanical piston injector is recommended for liquid pulsed plasma engine

When compared with all three suggested designs mechanical piston injector has several advantages over the gas compression and thermal injector. Because they carry out very simple mechanisms and they have got only the piston as a moving part they require very low power for operation. Gas compression injector also has same performance like mechanical piston but they require pumping device storage facilities for compressing gas and also for gas leakage from storage tank.

4.7.2 Designing Mechanical Piston Injector

From the deep research on each concept I have decided that the mechanical piston injector is the best injector for liquid pulsed plasma engine due to simplicity, easy to design and contains very simple mechanism.

The diagram below shows the mechanical piston injector, which has one moving part as a mechanical piston which is driven by high sensitive force generator, which has an ability to provide very accurate and small required force to the piston to inject the water. The propellant capacity of the injector is very small so on top the water container cylinder water propellant supply unit is connected. Soon after the propellant is fully consumed above, water supply unit will refill the propellant in very short time of period for next drive. The cylinder is connected to the water injection tube with small radius as shown in diagram, which will inject the propellant as the force is applied to the piston.

Designed mechanical piston injector (Figure 24)

4.7.3 Design calculation

Design Dimensions

Length of water injection tube: 1 cm Length of water containing cylinder: 5 cm Radius of water injection tube: 0.899 x 10-2 cm (calculated) Radius of water contains cylinder: 1 cm Temperature of water: 293K

Design constants Viscosity of water at 20’C: 1.002 x 10-3 Required mass bit: 2.5328 x10-7kg Time interval between each mass shot: 100s Density of water: 1000kgm-3

4.7.3 Calculation

Pressure points at injector (Figure 24)

From the previous calculation we found that the required mass bit per shot

Kg shot-1

Given that the time taken for each shot is 100µs

Kgs-1 =2.5328 x 10-3 Kgs-1

m =  x V m =  x A x X Mass is changing with respect to time, therefore X becomes (A)

Let assume that water is ejected with the speed of 100ms-1 from the injector = 100ms-1 Substituting the values in equation (A) gives,

A = 2.5328 x 10-8 m2

Where, is speed required to push the piston. A1 A2

3.1416  2.5328x10-8x100

0.8062cms-1

Consider the points P1, P2, and P3 on mechanical piston.

P = P2 – P1 = Constant P3 – P2 =

Apply the poisullies equation P = Where Q = Volume Flow Rate  = Viscosity of the Water = Length of the pipe R = Radius of pipe

But we know that P = F/A

So rearrange the equation as shown below

so

(B)

Required force to operate the mechanical piston to achieved the mass bit of 2.5328 x10-7kg

Fp1-p2 = 3.148mN

FP2-P3 Required, when piston is fully loaded = 1.269 x 10-3mN

Required initial force to push the piston F = Fp1-p2  + FP2-P3 F = 3.149 mN (it is very tiny different when comparing with Fp1-p2)

We can write equation A, as shown below (because L is changing with time)

=

= -2.046 x 10-4mNs-1

It means 2.046 x 10-4mN force is decreasing in each second. When we are comparing with Fp1-p2,   is very small. Therefore we can neglect this value.

F initial = (3.148+1.269 x10-3) mN          =3.149mN

F final = 3.148mN

Time required for drive the piston fully

Piston moves with a constant speed. Therefore acceleration is zero. So apply the linear motion Equation S = Ut + 1/2at2

Where S= distance of the water cylinder        (5cm) U = speed required to push the piston (0.8062cms-1) T = time taken for 1 drive

T=  =  = 6.2s

Finding the number of shots for one drive

100s = 1 shot 1s = 1/(100) shot 1s = 10,000 shot 6.2 s =6.2 10,000 shot =62,000 shot

The designed mechanical piston injector has an ability to deliver 62000 mass bit continuously per drive.

Conclusion

Liquid pulsed plasma thruster is an electromagnetic propulsion system and ideally comes from solid pulsed plasma thrusters. Both solid and liquid pulsed plasma thusters have same acceleration unit and similar function but they differ from the propellant element and their physical state. This project is about designing a propellant injector for pulsed plasma thruster. The project begins with basic introduction to the propulsion system and their classification. Mainly the spacecraft propulsion can be divided into two main subsystems they are chemical propulsion system and electric propulsion system. The chemical propulsion system is mainly used to take the payload from ground to space. Chemical propulsion system always creates high range of thrust with low specific impulse. The specific impulse is used to measure the efficiency of rocket performance. Electric propulsion system is mainly used in space mission operations and they create very low thrust with high specific impulse because they operate in vacuum condition and they are suitable for micro space craft for their altitude and directional controls. The electric propulsion system is mainly divided into three sub systems they are Electro static, Electro thermal and Electromagnetic propulsion system. Electro static system creates thrust by using Electro static energy to accelerate the ions. Electro thermal thrusters use Electro thermal energy to heat the propellant and pass through the nozzle to create the thrust. And Electro magnetic propulsion system uses electric energy to ionise the propellant convert into plasma stage and accelerate by applying both electric and magnetic field.

This project is mainly covering the Electro magnetic propulsion’s pulsed plasma thruster. They can be classified in three ways according to the propellant element’s physical stage. They are solid pulsed plasma thruster; Gas fed plasma thruster and liquid pulsed plasma thuster. The solid plasma thruster employs Teflon as propellant element. They ionise the propellant by applying high current and convert it into the plasma stage (plasma is a neutral gas with equal no of positive and negative ions) where they are accelerated to the high exhaust velocity by induced magnetic and electric field. Gas FED PPT also has same acceleration system like solid PPT. But here gaseous propellant is being used. Normally xenon or argon is used as propellant element. Gas FED PPT has high performance over the solid PPT. Their thrust level and specific impulse is much higher than solid PPT but they have some disadvantages as such as they require propellant storage unit and propellant leakage.

University of Tokyo recently developed a concept called liquid pulsed plasma thruster. They have used water as propellant element because they don’t contain carbon atom this carbon atom normally causes contamination to the spark plug. Therefore lifetime of the thruster can be increased. The main aim of this project is to design the propellant injector for liquid PPT. The University of Tokyo’s designed liquid injector is a silicon rubber valve type injector where they use simple silicon rubber valve mechanism to inject the water into the ionisation chamber. Earlier they tried to develop the injector with a pizzo electric device which did function in the atmospheric condition but did not function in vacuum condition. By using the equation (1) and I found the required mass bit for the liquid injector This is 2.5328 x 10-7kg. by considering that values I have introduce some concept.

In this project I have introduced three types of propellant injectors for liquid PPT. They are mechanical piston injector, gas compression injector and thermal injector. The gas compression injector uses helium gas to compress the liquid and eject into ionisation chamber. The thermal injector heats the solid form of water and converts it into liquid form where water will be sucked into the ionisation chamber due to the vacuum condition.

After the research on each injector according to the design requirements, mechanical piston satisfies all the requirements. They are very easy to use and have a very simple mechanism to inject the water; the system consists only of the piston as a moving part. I have developed my design process by considering the requied mass bit. The time taken for each mass bit is given as 100µs and by using that value I found the mass flow rate with respect to time.i made an assumption that the water leaves from the injector with velocity of 100ms-1. By considering that value I found the required radius for the injection tube is 0.89794 x 10-4m, this is a possible value for injector radius. And the speed required for moving the piston is also found which is 0.8062cms-1.there is also a possibility to move the piston with this velocity. Then the force required for pushing the piston to achieve required speed is calculated by using the poisollis equitation. The initial force required to push the piston is found as 3.149mN which changes in very negligible amount while the piston is moving along the cylinder, this is due to the change in pressure with respect to the time. Time taken for each drive is calculated by using the liner motion equation where we know the speed required to move the piston along the cylinder and length of the cylinder. By using these values I found that 6.2s is required for each drive of the piston. During that time the designed mechanical piston can eject the required mass bit of 62000 times per drive and it may require very short time to refill for the next drive.

References

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