Download AS301, IITM, Lecture 8 (22/08/2007)

Space Technology, Section-A, Module -4, Lecture-1
Class 10
ELECTRIC PROPULSION
Electrical and electromagnetic rockets differ fundamentally from
chemical rockets with respect to their performance limitations. Chemical rockets
are essentially energy limited, since the quantity of energy (per unit mass of
propellant) that can be released during combustion is limited by the fundamental
chemical behavior of propellant materials. Hence I sp of chemical rockets is limited
to 450 to 500 s. On the other hand, in electrical rockets a separate energy source
(eg. Nuclear or solar) is used, hence much higher propellant energy is possible.
The electric rocket engine is a device that converts electric power into a
forward-directed force or thrust by accelerating an ionized propellant (i.e.,
mercury, cesium, argon, or xenon) to a very high exhaust velocity. The concept
for an electric rocket is not new. In 1906 Dr. Robert Goddard (the famous
American rocket scientist) suggested that the exhaust velocity limit encountered
with chemical rocket propellants might be overcome if electrically charged
particles could be used as a rocket's reaction mass.
Advantages of Electrical Propulsion:
 Higher I sp than chemical rockets, usually ranges from 1000 to 20000 s.
 Electrostatic or Electromagnetic propellant acceleration can be achieved
without necessarily raising the fluid and solid temperature; hence there is
no upper limit to temperature of solid wall.
 There is no upper limit to the energy that can be added to the propellant.
 Life of satellite will be more.
Disadvantages of Electrical Propulsion:
 Lower thrust than chemical rockets, of the order of milli-Newton (mN)
hence cannot be used in launch phase.
 The rate of conversion of nuclear or solar energy to electrical energy and
thence to propellant kinetic energy is limited by the mass of conversion
equipment required.
 The mass of conversion equipment is a large portion of the total mass of
the vehicle, hence the electrical rocket is essentially power limited.
The basic electric propulsion system consists of three main components:
(1) some type of electric thruster that accelerates the ionized propellant; (2) a
suitable propellant that can be ionized and accelerated; and (3) a source of
electric power. The acceleration of electrically charged particles requires a large
quantity of electric power.
Electric propulsion is generally classified based on the input power source and
the operating principle of the thrusters (Figure 1).
Based on the input power source electric propulsion systems are classified as
nuclear electric propulsion (NEP) systems and solar electric propulsion (SEP)
systems. NEP use nuclear reactors in the spacecraft to generate power and SEP
use photovoltaic cells or solar thermal converters to generate power. Satellites
operating in a short range use SEP since the energy available from the Sun
decreases as [1/(distance)2] while NEP can be used even in deep space travel
as the nuclear power source is carried along with the space craft.
Electric propulsion
thrusters
Source of power
Solar
Operating principle
Nuclear
Electrothermal
Electrostatic
Electromagnetic
eg. Resitojet, Arcjet
eg. Ion thruster
eg. MPD thruster,
Pulsed plasma thruster
Figure 1. Classification of electric propulsion thrusters
Electric propulsion thrusters are classified into three types: Electrothermal,
Electrostatic and Electromagnetic based on the mechanism of using electrical
energy to generate thrust. In electrothermal thrusters, electrical energy is used to
heat the propellant and expand it thermodynamically in a nozzle to generate
thrust. The heating of the propellant gas is done either indirectly using a medium
to heat as in resistojet or directly by an electric arc discharge across the
propellant gas as in arcjet. In resistojets, the propellant gas is allowed to pass
over electrically heated solid surface (at around 3000 K) and expand in a nozzle
(Figure 2). Various heater configurations can be used in resistojets, such as
heater coils parallel to flow, transverse to the flow and honeycomb type parallel
flow passages etc. The heat losses are controlled by radiation shields,
regenerative cooling and thermal insulation. Hydrogen, ammonia and hydrazine
are the typical propellants employed. The input power ranges from 0.01-30 kW,
thrust ranges from 0.1-1 N and I sp values range from 250-1000 s. Thrust
produced can also be augmented in case of propellants such as hydrazine which
decomposes exothermically to produce high temperature products.
Radiation shield
Heating element
.
mp
Propellant inlet
Nozzle
Thermal insulation
Power source
Figure 2. Schematic of Resistojet
.
m
Arc (electrical
discharge)
Cathode
Constrictor
Anode
Power source
Figure 3. Schematic of Arcjet
Arcjets (Figure 3) use electrical arc discharge generated between inner cathode
and outer anode to heat the propellant gas passing through it. Since, arcjets heat
propellant directly; they have an advantage of higher thermal energy deposition
into the gas and also circumvent issues related to the maintenance of materials
(heating coils) operating at high temperatures as in resistojets. These simple
devices operate at power levels of 1-200 kW, thrust of 0.1-1 N and I sp range of
1000-2000 s. Generally, the electrical arc is unstable and its stabilization is a
critical factor to be considered in arcjet design. Constrictor passage and
tangential injection of propellant to induce a vortex flow are some of the methods
adopted to stabilize the arc.
electrons (negatively
charged)
Ions (positively
charged)
Ion source
Accelerating
grid electrode
Power source
Neutralizer
Figure 4. Schematic of Ion thruster (Electrostatic)
The electrostatic ion thruster eliminates the problems associated with the heating
of the propellant, wherein an electric field is used to accelerate the ions to
produce thrust. The thruster is made up of a source-accelerator-neutralizer
system (Figure 4) in which positive ions are produced and then made to
accelerate by an electrostatic field generated by anode and gridded cathode. To
maintain a zero net charge in the thruster, electrons from neutralizer are added to
the exiting positive ion beam. Being a developed concept in electric propulsion,
ion thrusters have very high range of I sp such as 5000-10000 s and thrusts less
than 0.5 N. Vapors of mercury and cesium and noble gases such as argon,
xenon and krypton are used as propellants.
High
voltage arc
discharge
Cathode
Anode
Current lines
.
mp
Lorentz force on plasma
Induced magnetic
field
Propellant
inlet
Anode
Accelerated plasma
Power source
Figure 5. Schematic of Self-field Magnetoplasmadynamic (MPD) thruster (Electromagnetic)
Magnetoplasmadynamic (MPD) thruster, generally being a steady flow
electromagnetic thruster, use electromagnetic body forces (Lorentz force)
generated in the ionized gas (plasma) by the interaction of current passing
through the plasma with the magnetic field (either induced by current passing
through the electrodes or externally applied). A schematic of a self-field MPD
thruster is shown in Figure 5. MPD thrusters are simple in construction but the
combination of complex physical processes involved have made these
electromagnetic devices difficult to implement. Still, MPD thrusters have the
potential of high thrust and high I sp combination required in deep space missions.
Ammonia, hydrogen and argon are the gaseous propellants employed. Lithium
has also been used as propellant as it reduces cathode erosion and gives high
thrust efficiency. The input power varies from few to several hundred kW with
thrust ranging of 0.1-10 N and I sp values ranging from 1000-5000 s.
Induced
magnetic field
Propellant
High voltage and
high current source
(capacitor/inductor)
Current
across the
plasma
Anode
Lorentz force accelerating the plasma layer
Cathode
Plasma formed
after high voltage
breakdown of cold
gas
Rail electrodes
Figure 6. Schematic of Pulsed Plasma Thruster (Electromagnetic)
Pulsed Plasma thruster (PPT) is an unsteady electromagnetic accelerator where
the highly ionized gas or plasma is produced, accelerated and ejected in a
pulsed manner. A high voltage and high current source across the electrodes
causes breakdown of the propellant, which creates an ionized column of gas.
The high current flowing across the electrodes induces a magnetic field which
generates electromagnetic body (Lorentz) force on the plasma column in
accordance with right hand grip rule.
Various configurations including two-
dimensional and coaxial configurations of PPT are present. Figure 6 shows a
schematic of 2D parallel rail configuration. Gases such as argon and solids such
as Teflon are used as propellants. Thrust values of around 0.01 N and I sp values
of 500-2000 s are possible. Being simple in design, PPTs can have long
operational periods (in years).
Analysis
A separate energy source (nuclear reactor), conversion device (to convert
nuclear/solar energy to electrical energy and then to propellant kinetic energy) is
required. This makes mass of power plant a large percentage of vehicle mass.
The initial electrical rocket mass is composed of M 0 = M L + M P + M s + M PP
Where M PP is the mass of the power plant
The sum of mass of power plant, M PP and space craft structure, M s increases
directly proportional to the power of the energy unit.
M PP + M s = αP where α is called specific mass.
Since electrical propulsion system is incapable of clearing earth’s gravity they are
carried by chemical rockets to orbit/deeper space where they can be used to
provide net positive thrust to the vehicle.
The electrical propulsion system are called power limited because of rate at
which energy from external source is supplied to the propellant is limited by the
mass available for the power system ( M PP < M 0 ).
The power of energy unit is proportional to the rate of kinetic energy expelled out
of the spacecraft.
m U e2
2
m U e2
or ηP =
where η is efficiency of thrust chamber
2
m U e2 TU e
=
Now T = m U e , so P =
2η
2η
P~
For a given thrust level, T , the power requirement, P , from the energy unit
increases as exhaust velocity is increased. Increase in power required results in
increased mass of power plant and structure.
M PP + M s = αP = α
Now T = m U e =
or M P =
TU e
~ U e (red curve in the figure below)
2η
M PU e
where tb is the burn time
tb
Tt b
1
~
(as U e decreases M P increases – black curve in figure below)
Ue Ue
At low U e the propellant mass fraction becomes high and at high U e power plant
and structural mass fraction is high. In both situations the payload mass is
reduced. Hence there is an optimal U e at which the payload fraction is maximum
(see figure below). Mathematical treatment of this optimization is discussed
below.
M0
M s + M pp + M P
ML
Mass
M s + M PP
MP
Ue
Figure 6. Various components of Launch vehicle mass as a function of exit velocity
Further Reading:
1. Rocket Propulsion Elements by G. P. Sutton and Oscar Biblarz (7th
Edition), Chapter#19, Electric Propulsion
2. Mechanics & Thermodynamics of Propulsion (2nd Ed.) by Hill & Peterson.
Chapter#10 section10.5
3. Physics of Electric Propulsion by Ribert G Jhan. Chapters 6, 7, 8, 9.
4. Space Propulsion Analysis and Design by R W Humble, G N Henry and
WJ Larson. 1st Edition (revised), Chapter#9.