Physics 101: Intro to Electronic Space Propulsion

Electric Propulsion (EP) represents a revolutionary approach to space propulsion, fundamentally differing from conventional chemical thrusters by utilizing electrical power to accelerate a propellant. This method significantly enhances propulsive performance and mass efficiency, making EP crucial for a wide range of modern and future space missions.

The underlying principle of electric propulsion is elegantly simple: solar power (minus efficiency losses in conversion) is used to ionize a propellant (e.g., xenon). These ions are then accelerated by a high-voltage electric field, converting their kinetic energy into a high-velocity exhaust. This high-velocity mass expulsion, in accordance with Newton's 3rd law, generates a change in momentum, which translates to a continuous thrust, propelling the spacecraft.

I. Fundamental Principles of Electric Propulsion

A. Core Concept:

Electric propulsion harnesses electrical and/or magnetic forces to accelerate a propellant, ejecting it at velocities up to twenty times faster than chemical thrusters. This high exhaust velocity leads to significantly increased mass efficiency, meaning less propellant is needed for a given change in velocity.

B. Key Advantages:

• High Mass Efficiency: Achieved by the extremely high exhaust velocities, leading to substantial propellant mass savings.
• High Specific Impulse ($I_{sp}$): EP thrusters typically boast specific impulses ranging from 1,000 to over 10,000 seconds, far exceeding chemical rockets (Isp​ values of 250-450 seconds). This translates to remarkable propellant efficiency and extended mission durations for a given amount of fuel.
• Long-Duration Thrust: While EP thrusters produce very low thrust (micro to milli-newton levels), they can operate continuously for months or even years. This allows spacecraft to accumulate significant velocity changes over time, enabling missions that are otherwise impractical with chemical propulsion.
• Energy-Limited, Not Propellant-Limited: Unlike chemical systems, which are constrained by the stored chemical energy in the propellant, EP is limited by the electrical power available onboard the spacecraft. This means that as long as power is supplied, the thruster can continue to operate, offering virtually unlimited energy for propulsion.
• Flexible Layout and High Automation: EP systems allow for flexible integration into spacecraft designs and can be highly automated for autonomous operation.

C. Limitations:

• Low Thrust and Acceleration: The primary drawback is the very low thrust, resulting in extremely low acceleration ($9.2\times 10^{-5}{\text{ m/s}}^2$ for a 1-ton satellite with 92 mN thrust). This makes EP unsuitable for launching spacecraft from planetary surfaces or for rapid orbital maneuvers.
• Power System Mass: The mass of the power generation and processing units can be substantial, directly correlating with the peak power required by the thruster. This can impact the overall spacecraft mass budget.
• System Complexity and Cost: Development and integration of EP systems can be more complex and expensive upfront compared to simpler chemical systems, though the long-term operational savings (due to reduced propellant and increased mission lifetime) often provide a significant return on investment.

II. Components of an Electric Propulsion System (EPS)

An Electric Propulsion System is typically composed of four main building blocks:

• Thruster Components: The core unit responsible for accelerating the propellant.
• Propellant Management System (Fluidic Components): Stores and precisely delivers the propellant to the thruster.
• Power Components: Includes the Power Processing Unit (PPU) which converts and conditions the electrical power from the spacecraft's power source to the specific voltages and currents required by the thruster.
• Pointing Mechanisms (Optional): Gimbals or other mechanisms to adjust the direction of thrust for maneuvering.

III. Types of Electric Propulsion Thrusters

Electric propulsion thrusters are broadly categorized by their primary method of accelerating the propellant:

A. Electrostatic Thrusters:

These thrusters accelerate ions using the Coulomb force along an electric field. Their principle relies on creating a cloud of positive ions by extracting electrons from neutral gas atoms. These ions are then accelerated by high-voltage electrodes. A neutralizer then re-injects electrons into the ion beam to prevent the spacecraft from accumulating a negative charge.

1. Gridded Ion Engines (GIEs):

   • Principle: Propellant atoms are ionized within a discharge chamber (e.g., by electron bombardment from a hollow cathode or by radiofrequency/microwave excitation). The resulting positive ions are then extracted and accelerated by a system of precisely aligned, multi-aperture grids (typically a positively biased screen grid and a negatively biased accelerator grid). The potential difference between these grids creates a powerful electric field that accelerates the ions to high velocities. The final ion energy is determined by the plasma potential, which is slightly greater than the screen grid's voltage.
   • Ionization Methods:
     • Electron Bombardment (Kaufman-type): Uses a hollow cathode to emit electrons, which are then accelerated to ionize the propellant (e.g., NASA's NSTAR, NEXT, T5, T6). Requires power supplies for the cathode, anode, and chamber.
     • Radiofrequency (RF) Excitation: Uses an oscillating electric field induced by an alternating electromagnet to create a self-sustaining discharge, omitting the need for a cathode or anode supplies (e.g., RIT 10, RIT 22, µN-RIT).
     • Microwave Excitation: Uses microwave heating to ionize the propellant (e.g., µ10, µ20). Similar to RF types, it can omit cathodes.
   • Extraction Grid Systems: Minor differences in grid geometry and materials (e.g., carbon-carbon composites, molybdenum alloys) can significantly impact operational lifetime, as grid erosion by charge-exchange ions is the primary life-limiting factor. Lifetimes of 20,000 hours or more can be achieved with thorough grid design.
   • Examples:
     • NASA's NSTAR (NASA Solar Technology Application Readiness): The first ion thruster to be used for primary propulsion on a deep-space mission (Deep Space 1), demonstrating long-duration operation.
     • NASA's Evolutionary Xenon Thruster (NEXT): An advanced ion propulsion system developed by NASA, offering higher power and efficiency than NSTAR.
     • ESA's T5/T6: European ion thrusters developed for various applications.
   • Applications: Deep space probes (e.g., NASA's Dawn mission to Vesta and Ceres), station-keeping for geostationary satellites, orbit raising, fine adjustments for scientific missions.

2. Hall Effect Thrusters (HETs):

   • Principle: HETs accelerate ions within an annular channel. Electrons, emitted from a cathode, are trapped by a radial magnetic field, creating a strong azimuthal (Hall) current. These trapped electrons ionize the incoming propellant gas. The resulting ions are then accelerated by an axial electric field. The magnetic field is strong enough to impede electron motion across the channel but weak enough to allow ions to be accelerated out of the thruster.
   • Examples: Stationary Plasma Thruster (SPT), Thruster with Anode Layer (TAL), Cylindrical Hall Thrusters, High Efficiency Multistage Plasma Thruster (HEMPT).
   • Applications: Widely used for orbit raising and station-keeping for commercial satellites, especially for large constellations (e.g., SpaceX's Starlink satellites primarily use Hall thrusters). HEMPTs, specifically, are a promising technology in Europe due to their high efficiency across a wide power range.

3. Field Emission Electric Propulsion (FEEP) Thrusters:

   • Principle: FEEP thrusters use a liquid metal propellant (e.g., indium or cesium). A strong electric field applied to the tip of a needle-like emitter ionizes the liquid metal through field emission, and the resulting ions are electrostatically accelerated to produce thrust.
   • Applications: Highly precise attitude control, fine positioning for sensitive scientific missions, and for very small satellites due to their compact size.

4. Colloid and Electrospray Thrusters:

   • Principle: These are similar to FEEP in that they accelerate charged droplets or ions from a liquid propellant. Electrospray thrusters typically produce finer sprays and primarily ions, while colloid thrusters might accelerate larger charged particles.
   • Applications: Micro-propulsion, attitude control for small satellites.

B. Electromagnetic Thrusters (Plasma Propulsion Engines):

These thrusters use the Lorentz force (F=q(E+v×B)) to accelerate all species (free electrons, positive ions, and negative ions) in the same direction, regardless of their electric charge. The electric field is not necessarily in the direction of the acceleration.

1. Pulsed Plasma Thrusters (PPTs):

   • Principle: A capacitor bank discharges a high-current pulse across a solid propellant (typically Teflon). This creates a brief, high-temperature plasma. The magnetic field generated by the high current interacts with the plasma, accelerating it via the Lorentz force.
   • Characteristics: Produce short, high-power pulses, enabling fine thrust control.
   • Applications: Small satellite propulsion, attitude control, and drag compensation for low Earth orbit.

2. Magnetoplasmadynamic (MPD) Thrusters:

   • Principle: A large electrical current is passed through a neutral gas, creating a plasma. The interaction of this current with its self-induced magnetic field, or an externally applied magnetic field, generates the Lorentz force that accelerates the plasma.
   • Characteristics: Capable of very high power (up to megawatts) and thrust levels (up to hundreds of Newtons) compared to other EP systems, but require substantial power and highly efficient power processing units.
   • Applications: Envisioned for high-power interplanetary missions, fast transit times for crewed missions.

3. Quad Confinement Thruster (QCT):

   • Principle: An electrode-less thruster that uses microwave energy to create and heat plasma, which is then accelerated by a magnetic nozzle. Its unique magnetic field configuration provides four regions of plasma confinement.
   • Characteristics: Eliminates electrode erosion, potentially leading to very long operational lifetimes.
   • Examples: Apollo Fusion's ACE and ACE Max thrusters (now part of Busek Co. Inc. after acquisition) are examples of QCTs.
     • ACE Datasheet
     • ACE Max Datasheet
   • Applications: Satellite constellations, orbit raising, and deep space missions where long life and high efficiency are critical.

C. Electrothermal Thrusters:

These systems use electromagnetic fields to generate a plasma or heat a propellant, which is then expanded through a nozzle to create thrust.

1. Resistojets:

   • Principle: Propellant gas is heated electrically by passing it over a resistively heated element. The hot gas is then expanded through a nozzle to produce thrust.
   • Characteristics: Higher specific impulse than cold gas thrusters, relatively simple.
   • Applications: Station-keeping, orbital maneuvers.

2. Arcjets:

   • Principle: Propellant gas passes through an electric arc, which superheats it to very high temperatures (thousands of Kelvin) before it is expanded through a nozzle.
   • Characteristics: Higher specific impulse than resistojets, but also higher power requirements.
   • Applications: Orbit raising and station-keeping.

3. Variable Specific Impulse Magnetoplasma Rocket (VASIMR):

   • Principle: Propellant is ionized into a plasma using radiofrequency waves, then heated further by additional radiofrequency waves, and finally directed and accelerated by a magnetic nozzle.
   • Characteristics: Offers the unique capability to vary specific impulse and thrust in real-time by adjusting power input, allowing for mission optimization (e.g., high thrust for rapid maneuvers, high Isp​ for long cruises).
   • Applications: Proposed for high-power, fast interplanetary missions, including human missions to Mars.

IV. Space Charge and its Significance

A. Definition:

Space charge is a concept where excess electric charge is treated as a continuous distribution over a region of space, rather than discrete point charges. It typically arises when charge carriers are emitted from a solid or when charged atoms/molecules are left behind.

B. Occurrence:

Space charge only occurs in dielectric media (including vacuum) because in conductive media, charge tends to be rapidly neutralized or screened. Both negative and positive space charges can form.

C. "Edison Effect":

This phenomenon, first observed by Thomas Edison in light bulb filaments when heated in a vacuum, describes the cloud of emitted electrons forming a negative space charge region around the filament.

D. Relevance to Electric Propulsion:

Space charge is a critical factor in the design and operation of ion thrusters.

• Neutralization: The positive space charge created by the accelerated ion beam would attract electrons from the spacecraft, causing the spacecraft to accumulate a net negative charge. This would, in turn, attract the expelled ions back toward the spacecraft, neutralizing the thrust. A neutralizer cathode is essential to inject low-energy electrons into the ion beam, ensuring that equal amounts of positive and negative charge are ejected and the exhaust remains electrically neutral.
• Beam Expansion and Interaction: Space charge forces within the ion beam cause it to expand, affecting the beam divergence and efficiency. Understanding and managing these effects are crucial for optimizing thruster performance and preventing interaction with spacecraft components.
• Operational Life: The negative voltage on the accelerator grid prevents electrons from the beam plasma outside the thruster from streaming back to the discharge plasma. Insufficient negative potential can lead to electron backstreaming, which is a common failure mode and can limit the operational life of ion thrusters.

V. Power Generation for Electric Propulsion

The efficiency and performance of EP systems are directly dependent on the available electrical power.

A. Solar Arrays:

• Dominant Source: Solar panels are the most common power source for Earth-orbiting satellites and inner solar system missions. They convert sunlight into electricity via the photovoltaic effect.
• Limitations: Power output decreases significantly with increasing distance from the Sun, making them less practical for deep space missions beyond the inner solar system.

B. Nuclear Power:

For missions far from the Sun or those requiring very high power, nuclear power sources are indispensable.

1. Radioisotope Thermoelectric Generators (RTGs):

   • Principle: RTGs generate electricity from the heat produced by the radioactive decay of radioisotopes (primarily Plutonium-238, with Americium-241 being explored as an alternative). Thermocouples convert this heat directly into electricity.
   • Characteristics: Highly reliable with no moving parts, capable of operating for decades. They produce relatively low power (tens to a few hundreds of watts) but cannot be varied or shut down once activated.
   • Applications: Essential for deep space probes (e.g., Voyager, Cassini, New Horizons, the Mars rovers Curiosity and Perseverance) where solar power is insufficient.
   • Cost & Supply: Plutonium-238 is expensive and in limited supply, driving research into more efficient RTG designs (e.g., Advanced Stirling Radioisotope Generator - ASRG) and alternative isotopes.

2. Fission Reactors (Space Reactors):

   • Principle: Utilize controlled nuclear fission of uranium to generate heat, which is then converted into electricity.
   • Characteristics: Capable of producing much higher power levels (kilowatts to megawatts) than RTGs.
   • Applications: Proposed for high-power electric propulsion missions to the outer solar system, enabling faster transit times for crewed missions to Mars, and for powering lunar or Martian bases. While historically utilized by Russia, new designs are being developed by the USA (e.g., Kilopower project) and other nations.

VI. Propellants for Electric Propulsion

The ideal propellant for an ion thruster is easy to ionize, has a high mass-to-ionization energy ratio, does not significantly erode the thruster components (for long life), and does not contaminate the spacecraft.

A. Noble Gases:

These are the most common propellants due to their inertness, ease of ionization, and relatively high atomic mass.

1. Xenon (Xe):

   • Advantages: High atomic mass, low ionization energy, chemically inert, high storage density as a liquid. These properties make it the most efficient noble gas propellant for many EP systems, especially Hall thrusters and gridded ion engines.
   • Disadvantages: Very high cost (e.g., several thousand Euros per kilogram).
   • Current Use: The primary propellant for a vast majority of commercial and scientific EP missions.

2. Krypton (Kr):

   • Advantages: Significantly lower cost than xenon, similar chemical characteristics, and can be stored as a liquid.
   • Disadvantages: Slightly higher ionization energy and lower atomic mass than xenon, leading to marginally lower performance (e.g., lower specific impulse and thrust-to-power ratio).
   • Current Use: Increasingly adopted for large satellite constellations (e.g., earlier versions of SpaceX's Starlink satellites) to reduce operational costs, accepting a slight performance trade-off.

3. Argon (Ar):

   • Advantages: The most abundant and cheapest noble gas (orders of magnitude less expensive than xenon and krypton), readily available.
   • Disadvantages: Significantly higher ionization energy and much lower atomic mass than xenon/krypton, resulting in lower efficiency and specific impulse for a given thruster design.
   • Current Use: Being explored and adopted for cost-sensitive applications, particularly for some large constellations (e.g., later versions of SpaceX's Starlink satellites are reported to use argon) where specific thruster designs can mitigate some of the performance losses.

B. Liquid Metals:

Used in FEEP thrusters.

• Examples: Indium, Cesium.
• Advantages: Very high specific impulse, extremely precise thrust control at very low levels.
• Disadvantages: Specialized applications, limited to very low thrust levels.

C. Other Propellants:

• Iodine (I2): Emerging as a promising alternative. It can be stored as a solid, simplifying tankage and reducing the need for high-pressure vessels. It can be vaporized and ionized.
  • Advantages: Storable as a solid (dense storage, no high-pressure tanks), lower cost than xenon, similar performance to xenon once ionized.
  • Disadvantages: Corrosive properties can pose material compatibility challenges.
  • Current Use: Actively being researched and demonstrated in orbit for small satellite applications.
• Bismuth (Bi): Another solid propellant candidate, similar to iodine in storage advantages.
• Conventional Propellants: In some electrothermal thrusters (resistojets, arcjets), conventional propellants like hydrazine can be used, heated electrically rather than combusted.

VII. Applications of Electric Propulsion

EP has become a cornerstone technology for a wide range of space missions:

• Low Earth Orbit (LEO):

  • Earth Observation & Earth Science: Precise orbit maintenance and maneuverability.
  • Constellations: Mass deployment and maintenance of large satellite networks (e.g., Starlink, OneWeb). EP significantly reduces launch mass and enables continuous orbital adjustments and deorbiting capabilities.
  • Continuous LEO Operations (Air-Drag Control): Counteracting atmospheric drag at lower altitudes to maintain orbit for extended periods.

• Medium Earth Orbit (MEO):

  • Navigation: Maintaining precise orbital positions for navigation satellite constellations (e.g., Galileo).

• Geostationary Orbit (GEO):

  • Telecommunications Satellites:
    • Electric Transfer from GTO to GEO: Instead of using a chemical kick stage, EP can slowly raise the satellite's orbit from a Geostationary Transfer Orbit (GTO) to GEO, saving significant propellant mass and allowing for a larger payload or a smaller, cheaper launch vehicle. This process typically takes weeks to months.
    • Station Keeping: Maintaining the satellite's precise position in geostationary orbit for its operational lifetime.

• Space Transportation:

  • Launcher Kick Stages: Providing additional propulsion after main launch vehicle stages separate.
  • Space Tugs: Dedicated spacecraft designed for orbital transfers, servicing, and logistics within space.

• Space Science and Exploration:

  • Interplanetary Cruise: Long-duration, high-efficiency propulsion for missions to other planets, asteroids, and deep space (e.g., NASA's Deep Space 1, ESA's SMART-1, NASA's Dawn mission). Continuous thrust over a long interval can achieve high velocities while consuming far less fuel than traditional chemical rockets.
  • Deep Space Probes: Enabling missions to the outer solar system where chemical propulsion would be prohibitively mass-intensive.
  • Fine Adjustments for Scientific Missions: High-precision attitude control and trajectory adjustments for sensitive scientific instruments.
  • Cargo Transport: Between propellant depots or for future lunar/Mars logistics.

• Long-Endurance Missions: EP's high specific impulse makes it ideal for missions requiring continuous thrust for extended periods.

• (Extreme) Fine and/or Highly Agile Attitude Control: For missions demanding very precise pointing capabilities.

VIII. Realistic Times and Costs

A. Mission Durations:

• Orbit Raising (GTO to GEO): Typically takes 4-12 months for EP-equipped satellites, compared to hours or days for chemically propelled satellites. This extended transfer time is a trade-off for significant propellant savings and thus lower launch mass.
• Interplanetary Missions: While initial acceleration is low, the continuous nature of EP allows for a much higher total change in velocity (delta-v) over time. This can lead to shorter overall trip times for missions requiring large delta-V, as the spacecraft carries much less propellant mass, freeing up mass for instruments or faster acceleration over a long duration.

B. Costs:

• System Development & Integration: Initial development and integration costs for EP systems can be higher than for conventional chemical systems due to complexity. However, the rapidly maturing market and increased demand are driving these costs down.
• Propellant Costs:
  • Xenon: Very high, on the order of $4,000 - $10,000 USD per kilogram depending on purity, quantity, and supplier.
  • Krypton: Significantly cheaper than xenon, typically in the range of $500 - $2,000 USD per kilogram.
  • Argon: The cheapest noble gas, often costing around $100 - $500 USD per kilogram.
  • Iodine: While still being commercialized, early estimates suggest costs significantly lower than xenon, potentially in the range of $100 - $500 USD per kilogram when fully scaled.
• Overall Mission Cost Savings: The primary financial benefit of EP comes from drastically reduced launch costs. By requiring less propellant mass, spacecraft can often use smaller, less expensive launch vehicles or fit more payload onto existing launchers. For instance, a GEO satellite with EP could save tens of millions of dollars in launch costs compared to a chemically propelled counterpart.
• Mission Lifetime Extension: EP's efficiency extends mission lifetimes (e.g., for station-keeping), leading to greater revenue generation for commercial satellites or more scientific data for research missions, providing a significant return on investment.

IX. Latest Advancements and Future Outlook

Electric propulsion technology is experiencing rapid growth and innovation.

• Increased Efficiency and Thrust-to-Power Ratio: Ongoing research focuses on maximizing the thrust produced per unit of electrical power, and improving the conversion of electrical energy into kinetic energy of the exhaust.
• Scalability: Development spans from micro-thrusters for CubeSats (e.g., for attitude control, deorbiting, and small delta-V maneuvers) to very high-power thrusters for future deep-space human exploration.
• Alternative Propellants: Iodine, bismuth, and even water are being actively investigated as potential replacements for xenon due to cost, storage advantages (solid-state), or abundance. This is a significant area of research to reduce operational costs for large constellations.
• Enhanced Lifetime and Reliability: Engineers are continuously working on improving the longevity of critical components like hollow cathodes and grids. Advanced materials (e.g., carbon-carbon composites for grids) and improved thruster designs are extending operational lifetimes to well over 20,000 hours, and often exceeding 30,000 hours in ground tests, allowing for multi-year missions. Preventing charge-exchange ion erosion of accelerator grids remains a key challenge and research area.
• Advanced Power Processing Units (PPUs): Development of more compact, lighter, and more efficient PPUs is crucial. These units convert and condition the spacecraft's main bus power to the precise voltages and currents required by the thruster, minimizing energy losses.
• Integrated Systems: Future EP systems aim for higher levels of integration, combining the thruster, propellant management system, and PPU into a single, compact module to reduce mass, volume, and integration complexity for spacecraft manufacturers.
• Advanced Diagnostics and Modeling: Sophisticated computational models and plasma diagnostics are being used to gain deeper insights into the complex physics within thrusters and their interaction with the spacecraft environment. This helps optimize designs, predict performance, and mitigate issues like plume impingement on sensitive spacecraft components.
• Hybrid Propulsion Architectures: Combining EP with chemical propulsion (e.g., chemical for initial high-thrust maneuvers, EP for long-duration cruise) is a growing trend to leverage the strengths of each.
• New Thruster Concepts: Research continues into novel thruster concepts beyond the established GIEs and HETs, seeking breakthroughs in efficiency, thrust density, or lifetime. Examples include advanced forms of electrodeless thrusters.

In Europe, significant developments have been carried out in all areas of electric propulsion over the last four decades. The most mature technologies (TRL 8-9) and promising for Europe include HET, GIE, and HEMPT. The selection of a specific EPS for a mission depends on a comprehensive system-level trade-off, considering factors like thrust capabilities, electrical power consumption, mission requirements, and flight heritage. The increasing flight heritage of EP systems is deeply influencing their selection for a wider array of missions.

References:

https://apollofusion.com/acemax.html
https://apollofusion.com/datasheets/Apollo_ACE_Datasheet-Jan_2021.pdf
https://apollofusion.com/datasheets/Apollo_ACE_Max_Datasheet-Jan_2021.pdf
https://www.esa.int/Enabling_Support/Space_Engineering_Technology/What_is_Electric_propulsion
https://www.esa.int/ESA_Multimedia/Images/2015/06/EPS_Main_building_blocks
https://apollofusion.com/ace.html
https://solarsystem.nasa.gov/missions/dawn/mission/faq/
https://en.wikipedia.org/wiki/File:Electrostatic_ion_thruster-en.svg
https://en.wikipedia.org/wiki/NASA_Solar_Technology_Application_Readiness
https://www.mathscinotes.com/2016/11/ion-propulsion-math/

https://en.wikipedia.org/wiki/Ion_thruster

https://en.wikipedia.org/wiki/Spacecraft_propulsion

https://en.wikipedia.org/wiki/Variable_Specific_Impulse_Magnetoplasma_Rocket

https://www.youtube.com/watch?v=XIqDHsmsbPc