electric propulsion systems

7/29/2019 electric propulsion systems

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Propulsion in a broad sense is the act of changing the motion of a body where a reaction

force is imparted to a device by the momentum of ejected matter.Propulsion mechanisms provide a force that moves bodies that are initially

at rest, changes a velocity, or overcomes retarding forces when a body is

propelled through a medium.

The propellants, which are the working substance of rocket engines, constitute

the fluid that undergoes chemical and thermodynamic changes.

SOLID: Highest power outputs but easy to control . Lower burn times can be achieved

LIQUID: Higher power outputs ,relatively difficult to control in comparison with solid

propellants. Higher burn times

EXAMPLE: Hydroxyl Ammonium Nitrate (NH2OH*NO3)

COLD GAS PROPELLANTS: They are simple ,low cost, used with pressurized feed systems, used for pulsing

operations, and for low thrust and low total impulse.

EXAMPLE: Liquid Oxygen+Liquid hydrogen

EXAMPLE:Nitrocellulose+Nitroglycerine


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CHEMICAL PROPULSION

The energy from a high-pressure combustion reaction of propellant chemicals,

usually a fuel and an oxidizing chemical, permits the heating of reaction productgases to very high temperatures (2500 to 4100C or 4500 to 7400F).

These gases subsequently are expanded in a nozzle and accelerated to high velocities (1800 to 4300 m/sec

or 5900 to 14,100 ft/sec). Since these gas temperatures are about twice the melting point of steel, it is

necessary to cool or insulate all the surfaces that are exposed to the hot gases

.

Advantages:

a) Very high values of thrust. Therefore they are used in the initial launch where the

gravitational forces are very high.

b) Very low burn time required for attaining definite acceleration.

c) Temperature of the thrust chamber can be controlled easily

Disadvantages:

a) Large storage tank for fuel required 40% of the mass of the satellite is propellant .

b) Only the energy contained in the propellant can be used for acceleration thus only limited

velocity is obtainable (about 4.5km/sec)

Application:

1. First stage and its upper stage propulsion systems add momentum during launch and ascent.2.Orbit injection or transferring from one orbit to another.


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Solar thermal rocket, uses large diameter optic mirrors to concentrate the sun's radiation (e.g.

lightweight precise parabolic mirrors or Fresnel lenses) onto a receiver or optical cavity.

SOLAR PROPULSION

Fig 1.2 :. Simplified schematic diagram of a solar thermal rocket concept

Advantages:

a) High efficiency,

b)No need of oxygen tanks so reduced weights

c) Contamination is negligible,,

Disadvantages: a)

Since large lightweight optical elements cannot withstand drag forces without deformation, the

optical systems are deployed outside the atmosphere.

Application:They can be used in station keeping


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NUCLEAR PROPULSION

Various nuclear energy sources have been investigated for delivering heat to a working fluid,

usually liquid hydrogen, which subsequently can be expanded in a nozzle and thus accelerated

to high ejection velocities (6000 to 10,000 m/sec). They are basically extensions of liquid

propellant rocket engines. The heating of the fuel is accomplished by energy derived fromtransformations within the nuclei of atoms.

Fig 1.1 : The principle of nuclear thermal propulsion.

Advantages: a) Very high thrust values (about 40000N)

b) Considerably high exhaust velocities (6000 to 10,000 m/sec).

Disadvantages: a) The thrust chamber temperatures are extremely high and are difficult to cool

b) There is a possibility of radiation leakage.

C) Low efficiencyApplication:

Flight has not been confirmed yet but they can be used in first and upper stage propulsion and also for orbit injection.


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ELECTRIC PROPULSION

Electric rocket propulsion devices use electrical energy for heating and/or directly ejecting

propellant, utilizing an energy source that is independent of the propellant itself.

In all electric propulsion the source of the electric power (nuclear, solar radiation receivers, or batteries) is

physically separate from the mechanism that produces the thrust.Electric rocket propulsion devices useelectrical energy for heating and/or directly ejecting propellant, utilizing an energy source that is

independent of the propellant itself. The thrust usually is low, typically 0.005 to 1 N. In order to allow a

significant increase in the vehicle velocity, it is necessary to apply the low thrust and thus a small

acceleration for a long time.

Advantages:a)Very high efficiencies(about 90%).

b)Only 0.3% of the rocket is due to the fuel.

c)Very high exhaust velocities

d)Improve the rockets mass ratio significantly .

Disadvantages:

a)Very low thrust 0-1N(although this low thrust has been put to use in various types of maneuversefficiently)

b)Very high Burn times

c) This type of propulsion has been handicapped by heavy and inefficient power sources

Application:

a) Pitch correction of rocket

b)These propulsion systems have beenextensively used in station keeping


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1.ELECTRIC PROPULSION SYSTEMS HAVE A VERY HIGH EFFICIENCY(about

90%).Whereas Chemical propulsion systems have a maximum efficiency of

about 60%.

2.A very large Exhaust Velocity can be easily obtained by using Electrical

Propulsion. Which is a major challenge in case of Chemical Propulsion as the

exhaust velocity of only 4.5km/sec obtainable. This is possible due to

formidable increase in the mass ratio. So for deep space missions Electrical

propulsion systems is most suitable.

3.The effective thrust produced by an Electric propulsion system is very low.

So it cannot be used in order to overcome high gravitational forces ,and

hence can be employed only outside orbits of planets. But certain specifictasks performed by spacecraft such as docking, space vehicle manuvering

etc are only possible due to low thrust of electric propulsion systems.

4.This propulsion systems can be used in applications where the burn time is

very high .So they are employed where the mission time last for more than

10 years.


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ENGINE TYPE

Propulsive

efficiency

INPUT

CURRENT

Is (A)

EXHAUST

VELOCITYPOWER INPUT

Chemical rocket

0.50 30 2940 29.4 kW

Nuclear fission 0.50 80 7840 78.7kW

Solar 0.50 60 5880 58.8Kw

Ion electrostatic 0.90 200 10,600 195.9Kw

Here we can see that electric propulsion systems have a very high efficiency as compared to

other system. Moreover the exhaust velocities are very high so velocity increment is very high

which is primary characteristic for deep space missions


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The basic subsystems of a typical electric propulsion thruster are: (1) An energy source

(2)Conversion devices to transform this energy into electrical form at the proper voltage, frequency, pulse

rate, and current suitable for the Electrical propulsion System .

3)Electric Propulsion system .

1.Energy sources:

Batteries

Fuel Cells

Solar Cell Arrays

Long-Duration High-Output Dynamic System

2.Power-Conditioning Equipment

Modern conditioning equipment contains all the internal logic required to start, safely operate, and stop the

thruster; it is controlled by on-off commands sent by the spacecraft-vehicles and rockets control processor

.In fact, advances in solid-state electronic pulse circuits together with lighter, more efficient, and higher

temperature power-conditioning hardware are an area of great interest to the implementation of electric

propulsion units. The efficiency of the equipment tends to be high, about 90% or more, but the heat

generated is at a low temperature and must removed to maintain the required moderately lowtemperatures of operation


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Types of Electric Propulsion systems

CLASSIFICATION

1. Electro thermal. Propellant is heated electrically and expanded thermodynamically;

i.e., the gas is accelerated to supersonic speeds through a nozzle, as in the chemical

rocket.

2. Electromagnetic. Acceleration is achieved by the interaction of electric and magnetic

fields within a plasma. Moderately dense plasmas are high temperature or nonequilibrium gases, electrically neutral and reasonably good conductors of electricity.


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Electro thermal thrusters

The basic electrothermal thruster, or `resisto-jet', consists of a nozzle with a high

expansion ratio, connected to a chamber in which the propellant is heated by a hot

wire through which passes an electric current. This type of electric thruster uses the

same thermodynamic effects to generate a high-velocity exhaust stream as does a

chemical rocket.

The propellant is heated by flowing over an ohmically heated refractory-metal surface, such as

(1) coils of heated wire,

(2) through heated hollow tubes,

(3) over heated knife blades, and

(4) over heated cylinders

Fig: The Schematic of Resistojet Propulsion System


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Power requirements range between 1 W and several kilowatts; a broad range of terminal

voltages, AC or DC, can be designed for, and there are no special requirements for power

conditioning

The system using liquid propellants has the advantage of being compact and the catalytic

decomposition preheats the mixed gases to about 700C (1400F) prior to their being heated

electrically to an even higher temperature; this reduces the required electric power while taking

advantage of a well-proven space chemical propulsion concept

Thruster efficiencies of resistojets range between 65 and 85%, depending on the propellant

and the exhaust gas temperature,

They are most attractive for low to modest levels of mission velocity increments, where

power limit ,thrusting times, and plume effects are mission drivers.


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Arcjet

Fig 2.3 :The schematic of an arc jet

Arcjet overcomes the gas temperature limitations of the resistojet by the use of an

electric arc for direct heating of the propellant stream to temperatures much higher

than the wall temperatures. The arc stretches between the tip of a central cathode and

an anode, which is part of the coaxial nozzle that accelerates the heated propellant

For electrical (but not thermal) insulation, boron nitride has been used effectively. In reality, arcs are

highly filamentary and tend to heat only a small portion of the flowing gas unless the throat dimension issufficiently

The conduction of electricity through a gas requires that a certain level of ionization be

present. This ionization must be obtained from an electrical discharge, i.e., the breakdown

of the cold propellant


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Steady thrust 222-258 mN

Feed pressure 185-330pa

Power control unit (PCU) input 4.4 kW

PCU efficiency 93%

Dimensions; Arcjet 237 x 125 x 91

PCU -632x361 x 109

Mass: Arcjet and cable 6.3 kg

PCU 15.8 kg

Specific performance characteristics of arc jets

Arcjets, however, are potentially more scalable to large thrust levels than other electric propulsionsystems Arcjets have another disadvantage in that the required power processing units are somewhat

more complex than those for resistojets, due to the complexity of arc phenomena.

The life of an arcjet can be severely limited by local electrode erosion and vaporization, which is

specifically due to action of the arc attachment point and of the high operating temperatures in general.


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Electromagnetic thrusters:

Electrostatic thrusters rely on Coulomb forces to accelerate a propellant composed of non-

neutral charged particles. They can operate only in a near vacuum. The electric force depends

only on the charge, and all charged particles must be of the same "sign" if they are to move in

the same direction. Electrons are easy to produce and are readily accelerated, but they are asextremely light in mass as to be impractical for electric propulsion.

Thus, the thrust per unit area that can be imparted to such an electron flow remains

negligible even when the effective exhaust velocity or specific impulse gets to be very high.

Accordingly, electrostatic thrusters use charged heavy-molecular-mass atoms as positive ions

(a proton is 1840 times heavier than the electron and a typical ion of interest containshundreds of protons)

In terms of power sources and transmission equipment, the use of the heavier particles contributes

to more desirable characteristics for electrostatic thrusters--for example, high voltages and low

currents.


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FIG 2.4 Simplified schematic diagram of an electron bombardment ion thruster,

The electrostatic (or Coulomb) force and the electromagnetic (or Lorentz) force can be used to accelerate a

suitable propellant to speeds

The microscopic vector force Fe on a singly charged particle can

be written as

Fe = eE + e(Ve B)

Where e is the electron charge magnitude, E the electric field vector, Ve the velocity of the charged particle, and B the magnetic

field vector. The exhaust velocity is a function of the voltage Vacc imposed across the accelerating chamber or grids, the mass of

the charged particle and its electrical charge e


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Ionization Schemes.

To a great extent, the ionization chamber is responsible for most of the size, mass, and perhaps efficiency of these

devices.Ionization of a gas by electron bombardment is a well-established technology. Electrons are emitted from athermionic (hot) cathode or the more efficient hollow cathode and are forced to interact with the gaseous propellant flow in

a suitable ionization chamber. The chamber pressures are low, typically 10 -3 torr or 0.134 Pa

Emitted electrons are attracted toward the cylindrical anode but are forced by the axial magnetic field to spiral in the chamber,

causing numerous collisions with propellant atoms which lead to ionization. A radial electric field removes the electrons from the

chamber and an axial electric field moves the ions toward the accelerator grids.

Electric Bombardment

Radiofrequency

Most high power hence high thrust systems use electrodes, of one kind or another, to generate the current in the gas that

provides the ions and hence the thrust. These always erode in the discharge, being worse for high power systems. Several

attempts have been made to increase the power input by using radiowaves to provide the internal energy source . The process

here begins with the gas, hydrogen and helium mixed, being exposed to the electromagnetic radiation from an RF antenna that

ionises most of the atoms. It then passes into a cavity with strong magnetic fields, where it is excited by a high-power

RFgenerator resonance here gives avery high efficiency of power transfer from the RF field to the gas.

Two aspects of the exhaust beam are non-thrust producing: one is the aforementioned ionization energycontained in the beam and

the other one is any vector divergence present which results in beam spreading. Beam spreading, or the radial velocity component of

beams, can result from causes both upstream and downstream of the exit

Proper neutralization of the beam reduces this spreading, allowing nearly axial velocities.

H th El t i P l i t i t t d


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How the Electric Propulsion system is started:

Step1:The sensing element (Transducers)for determining the attitude, velocity, and position of the vehicle with respect to areference direction at any one time convert the parameter in electrical signals and sends it to a microcontroller

Step2:The microcontroller analyses the input and checks it with the reference input received from ground station by microwaves or

pre programmed value of the parameter.

Step3:The Power conditioning equipment converts the power from source to desired values.and triggers the propulsion system .

Step4:The servo orifice valve opens and the propellant reaches the Electric Propulsion system.

The input to an electro-hydraulic (EH) servo valve is typically a current or a differential current that powers an electromagnetic torquemotor. The differential current i istypically supplied by an amplifier to avoid excess loading of the interface to the computer

or controller. In the simplest)form, the torque motor moves a spool valve as shown below. The spool valve allows the hydraulic fluid

to pass from the supply to the return across two variable metering orifices with a controlled flow rate QL. If the spool is shifted in the

other direction the direction of f low will reverse.

Fig :Schematic of an Electro Hydraulic Valve

Step5:The electric Propulsion system gets fully operational and thrusts the vehicle in the desired direction


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FIG:Space craft with thruster for Velocity vector adjustment

Velocity vector adjustment and minor in-flight correction maneuvers are usually performed with low

thrust, short duration and intermittent (pulsing) operations, using a Propulsion control system with multiple

small liquid propellant thrusters, both for translation and rotation. The vernier rockets on a ballistic missile

are used to accurately calibrate the terminal velocity vector for improved target accuracy. The reaction

control rocket systems in a space launch vehicle will allow accurate orbit injection adjustment maneuvers

after it is placed into orbit by another, less accurate propulsion system. Mid-course guidance-directedcorrection maneuvers for the trajectories of deep space vehicles fall also into this category.


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Deep Space

Lunarand interplanetarymissions include circumnavigation, landing, and

return flights to the moon, Venus, Mars, and other planets. The energy necessaryto escape from earth can be calculated as M

. It is 6.26 x 107j/kg, which is more than that required for a satellite.Such high Energy are only possible because Electric Propulsion System has the

exhaust velocities in the range of 10000m/sec

Low Earth orbit to geostationary orbit raising

The required velocity increment for an elliptical transfer to geo stationary altitude and

circularisation of the orbit is theoretically 4.2km/s. To this should be added the

gravity losses associated with the continuous burning of the electric propulsion

system, which is of the order 1km/s, producing a total of 5.2km/s for this mission.

This is 10 times the station keeping requirement. It should also be carried out in a

reasonable length of time, so may be expected to be more demanding of the

thrusters.

As a comparison chemical thruster we can consider a bi-propellant engine using

UDMH and nitrogen tetroxide. It has an exhaust velocity of 3,160m/s in vacuum

requiring a propellant fraction of 80% For the ion engine

the propellant fraction is 16%, and so the use of electric propulsion can potentially

save about 80% of the propellant. Note that the higher exhaust velocity of the ion

engine has a much bigger effect for this mission..


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Station keeping

Here a typical satellite will carry sufficient control propellant to produce a totalvelocity increment of 500m/s for station keeping. If this is provided by hydrazine

thrusters with an exhaust velocity of 1,000m/s, the ratio of propellant mass to total

vehicle mass is 0.39; 40% of the mass of the satellite is propellant. If an Ion thruster

with an exhaust velocity of 10km/s is used, then the propellant is only 3.3% of the

total mass of the satellite, and more than 90% of the propellant mass can therefore

be saved. This reduces launch costs (at $20,00030,000 per kg), or allows a largerSatellite to be launched for the same cost.

electric propulsion systems

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