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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.