Pulsed plasma thruster

Schematic layout of a Pulsed Plasma Thruster

Most PPTs use a solid material (normally PTFE, more commonly known as Teflon) for propellant, although very few use liquid or gaseous propellants. The first stage in PPT operation involves an arc of electricity passing through the fuel, causing ablation and sublimation of the fuel. The heat generated by this arc causes the resultant gas to turn into plasma, thereby creating a charged gas cloud. Due to the force of the ablation, the plasma is propelled at low speed between two charged plates (an anode and cathode). Since the plasma is charged, the fuel effectively completes the circuit between the two plates, allowing a current to flow through the plasma. This flow of electrons generates a strong electromagnetic field which then exerts a Lorentz force on the plasma, accelerating the plasma out of the PPT exhaust at high velocity.[1] Its mode of operation is similar to a railgun. The pulsing occurs due to the time needed to recharge the plates following each burst of fuel, and the time between each arc. The frequency of pulsing is normally very high and so it generates an almost continuous and smooth thrust. While the thrust is very low, a PPT can operate continuously for extended periods of time, yielding a large final speed.

The energy used in each pulse is stored in a capacitor.[3] By varying the time between each capacitor discharge, the thrust and power draw of the PPT can be varied allowing versatile use of the system.[2]

The equation for the change in velocity of a spacecraft is given by the rocket equation as follows:

${\displaystyle \Delta v=v_{\text{e}}\ln {\frac {m_{0}}{m_{1}}}}$

where:

${\displaystyle \Delta v\ }$ is delta-v - the maximum change of speed of the vehicle (with no external forces acting),

${\displaystyle v_{\text{e}}}$ is the effective exhaust velocity (${\displaystyle v_{\text{e}}=I_{\text{sp}}\cdot g_{0}}$ where ${\displaystyle I_{\text{sp}}}$ is the specific impulse expressed as a time period and ${\displaystyle g_{0}}$ is standard gravity),

${\displaystyle \ln }$ refers to the natural logarithm function,

${\displaystyle m_{0}}$ is the initial total mass, including propellant,

${\displaystyle m_{1}}$ is the final total mass.

PPTs have much higher exhaust velocities than chemical propulsion engines, but have a much smaller fuel flow rate. From the Tsiolkovsky equation stated above, this results in a proportionally higher final velocity of the propelled craft. The exhaust velocity of a PPT is of the order of tens of km/s while conventional chemical propulsion generates thermal velocities in the range of 2–4.5 km/s. Due to this lower thermal velocity, chemical propulsion units become exponentially less effective at higher vehicle velocities, necessitating the use of electric spacecraft propulsion such as PPTs. It is therefore advantageous to use an electric propulsion system such as a PPT to generate high interplanetary speeds in the range 20–70 km/s.

NASA's research PPT (flown in 2000) achieved an exhaust velocity of 13,700 m/s, generated a thrust of 860 µN, and consumed 70 W of electrical power.[1]

PPTs are very robust due to their inherently simple design (relative to other electric spacecraft propulsion techniques). As an electric propulsion system, PPTs benefit from reduced fuel consumption compared to traditional chemical rockets, reducing launch mass and therefore launch costs, as well as high specific impulse improving performance.[1]

However, due to energy losses caused by late time ablation and rapid conductive heat transfer from the propellant to the rest of the spacecraft, propulsive efficiency (kinetic energy of exhaust / total energy used) is very low compared to other forms of electric propulsion, at around just 10%.

PPTs are well-suited to uses on relatively small spacecraft with a mass of less than 100 kg (particularly CubeSats) for roles such as attitude control, station keeping, de-orbiting manoeuvres and deep space exploration. Using PPTs could double the life-span of these small satellite missions without significantly increasing complexity or cost due to the inherent simplicity and relatively low cost nature of PPTs.[3]

The first use of PPTs was on the Soviet Zond 2 space probe on 30 November 1964.

A PPT was flown by NASA in November, 2000, as a flight experiment on the Earth Observing-1 spacecraft. The thrusters successfully demonstrated the ability to perform roll control on the spacecraft and demonstrated that the electromagnetic interference from the pulsed plasma did not affect other spacecraft systems.[1] Pulsed plasma thrusters are also an avenue of research used by universities for starting experiments with electric propulsion due to the relative simplicity and lower costs involved with PPTs as opposed to other forms of electric propulsion such as Hall-effect ion thrusters.[2]

• Vacuum arc thruster

• "Design of a High-energy, Two-stage Pulsed Plasma Thruster". Princeton University. Archived from the original on 2017-08-08. Retrieved 2020-06-27.
• "EO1 Pulsed Plasma Thruster" (PDF). Goddard Space Flight Center. Archived from the original (PDF) on 2011-07-16. Retrieved 2020-06-27.
• Ephraim Chen. "Gas-Fed Pulsed Plasma Thrusters: From Sparks to Laser Initiation" (PDF). Princeton University. Archived (PDF) from the original on 2022-10-09. Retrieved 2020-06-27.
• Michael Bretti. "AIS-gPPT3-1C Single-Channel Gridded Pulsed Plasma Thruster". Applied Ion Systems. Retrieved 2020-06-27.