Ablative Pulsed Plasma Thruster Research Articles

Results of the qualification test campaign of a Pulsed Plasma Thruster for Cubesat Propulsion (PPTCUP)

Pulsed Plasma Thruster for Cubesat Propulsion (PPTCUP) is an ablative pulsed plasma thruster designed with the aim of providing translational and orbital control to Cubesat platforms. The qualification model presented in this paper has been developed by Mars Space Ltd, Clyde Space Ltd and the University of Southampton to produce a versatile “stand-alone” module that can be bolted on the Cubesat structure, allowing the orbital control along the X or Y-axis of the satellite. An extensive and complete test campaign to qualify the unit for space flight, which includes electromagnetic compatibility (EMC) characterization, thermal cycling and mechanical tests, has been performed according to the NASA GEVS procedures. PPTCUP is characterized by an averaged specific impulse of 655±58s and a deliverable total impulse of 48.2±4.2Ns. Finally, it has been found that the unit is compliant with the EMC requirements and can successfully withstand the thermal and mechanical loads typical of a Cubesat space mission.

Application of Mechanical Probes for Evaluation of Plasma Acceleration in Ablative PPT

In pulsed plasma thrusters (PPTs), a highly ionized plasma is created and accelerated by means of a pulsed triggered vacuum gap discharge with concurrently occurring ablation, ionization, and Lorentz-force-based acceleration of propellant. The plasma flow behavior during the PPT discharge is, therefore, highly dependent on time, space, and operation parameters. Experimental work is seen as the necessary key to further understand the processes involved and to determine future research needs. The application of mechanical probes to measure the Mach number of the plasma flow optically is discussed in this paper. Values of 1.6–3.2 were determined for a variation in time, space, and discharge voltage. An inclination of the PPT plasma flow toward the cathode was observed, which concurs with the previous research results.

Guest Editorial Introduction to the Special Issue on Plasma Propulsion

This Special Issue is dedicated to the physics, technology, and application of plasma propulsion for spacecraft. The field of plasma propulsion includes a broad variety of thrusters that can achieve high propellant exhaust velocity, thereby offering a large mass savings for space vehicles compared to chemical (combustion) rockets. These thrusters are broadly categorized by their propellant acceleration mechanism into three groups: 1) electrothermal; 2) electrostatic; and 3) electromagnetic <xref ref-type="bibr" rid="ref1" xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">[1]</xref> . Research into plasma propulsion dates back several decades, with a first application in space in 1964 on the Soviet Zond-2, which used an ablative pulsed plasma thruster to control the spacecraft orientation. Today, plasma propulsion is a very rapidly growing area of plasma science and technology, with over 250 operational spacecraft employing plasma propulsion in a variety of applications. Many new plasma thrusters have been recently developed, including numerous successful attempts to scale previously known systems to lower and higher power levels, ranging from a few watts to over 100 kW <xref ref-type="bibr" rid="ref2" xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">[2]</xref> , <xref ref-type="bibr" rid="ref3" xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">[3]</xref> . As space exploration shifts toward small and efficient satellites, or micro and nanosatellites <xref ref-type="bibr" rid="ref4" xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">[4]</xref> , <xref ref-type="bibr" rid="ref5" xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">[5]</xref> , there are many near-future space missions involving science, military, and commercial payloads utilizing micro and nanosatellite platforms. These platforms require very small levels of thrust for very fine attitude control, high resolution. Earth imaging and astronomy, as well as very precise positioning requirements and other very precise positioning requirements for spacecraft performing formation flying and interferometry missions. Miniaturized propulsion systems are required to satisfy these emerging needs for both the low-thrust missions and the propulsion on small-sized spacecraft.

Physicomathematical model of plasma acceleration in an ablative pulsed plasma thruster

A new comparatively simple quasi-one-dimensional physicomathematical model of plasma acceleration in an ablative pulsed plasma thruster with a capacitive energy storage is proposed. In spite of its simplicity, the model adequately reflects the main physical processes occurring in the thruster channel in the course of plasma blob acceleration: the blob dynamics, plasma radiation, absorption of radiation by the Teflon channel walls, ablation of the wall material, and plasma ionization. The results of computer simulations agree well with the experimental results.

High precision micro-impulse measurements for micro-thrusters based on torsional pendulum and sympathetic resonance techniques

A sympathetic resonance theory is analyzed and applied in a newly developed torsional pendulum to measure the micro-impulse produced by a μN s-class ablative pulsed plasma thruster. According to theoretical analysis on the dynamical behaviors of a torsional pendulum, the resonance amplification effect of micro-signals is presented. In addition, a new micro-impulse measurement method based on sympathetic resonance theory is proposed as an improvement of the original single pulse measurement method. In contrast with the single pulse measurement method, the advantages of sympathetic resonance method are significant. First, because of the magnification of vibration signals due to resonance processes, measurement precision for the sympathetic resonance method becomes higher especially in reducing reading error. With an increase in peak number, the relative errors induced by readout of voltage signals decrease to approximately ±1.9% for the sympathetic resonance mode, whereas the relative error in single pulse mode is estimated as ±13.4%. Besides, by using the resonance amplification effect the sympathetic resonance method makes it possible to measure an extremely low-impulse beyond the resolution of a thrust stand without redesigning or purchasing a new one. Moreover, because of the simple operational principle and structure the sympathetic resonance method is much more convenient and inexpensive to be implemented than other high-precision methods. Finally, the sympathetic resonance measurement method can also be applied in other thrust stands to improve further the ability to measure the low-impulse bits.

Study on electron temperature in an ablative pulsed plasma thruster by optical emission spectroscopy

The electron temperature obtained by optical emission spectroscopy was studied in an ablative pulsed plasma thruster. The validity of partial local thermodynamic equilibrium criteria for the correct spectroscopic determination of the electron temperature can be proved and the spectral absorption of the plasma was calculated and ignored. It is shown that the existence of two distinguishable electron groups is found. The variation trends of the part data of the experiment are in agreement with those of Gatsonis and Markusic under some conditions, respectively. Combined with the discharge current, the variation of the electron temperature and the transferring process of electron energy were discussed.

Characteristics of plasma properties in an ablative pulsed plasma thruster

Pulsed plasma thrusters are electric space propulsion devices which create a highly transient plasma bulk in a short-time arc discharge that is expelled to create thrust. The transitional character and the dependency on the discharge properties are yet to be elucidated. In this study, optical emission spectroscopy and Mach-Zehnder interferometry are applied to investigate the plasma properties in variation of time, space, and discharge energy. Electron temperature, electron density, and Knudsen numbers are derived for the plasma bulk and discussed. Temperatures were found to be in the order of 1.7 to 3.1 eV, whereas electron densities showed maximum values of more than 1017 cm−3. Both values showed strong dependency on the discharge voltage and were typically higher closer to the electrodes. Capacitance and time showed less influence. Knudsen numbers were derived to be in the order of 10−3−10−2, thus, indicating a continuum flow behavior in the main plasma bulk.

A new stage in the development of ablative pulsed plasma thrusters at the RIAME

This paper presents an overview of the works on ablative pulsed plasma thrusters (APPTs) carried out at the Research Institute of Applied Mechanics and Electrodynamics (RIAME). The main features of next generation thrusters developed at the RIAME in the 2000s are discussed together with the optimization criteria for APPTs intended for use in correction propulsion systems of small spacecraft, e.g., MKA-FKI (developed by the Lavochkin Association) and Soyuz-Sat-O (developed by Maksimov Space Systems Research Institute and the Production Corporation Polyot).

Thrust efficiency optimization of the pulsed plasma thruster SIMP-LEX

The effect of electric parameters on the thrust efficiency of an ablative pulsed plasma thruster was studied. Analytically, it was shown that a higher efficiency can be obtained by increasing energy of a bank of capacitors. This can be achieved by changing the inductance per distance of the plasma sheet, or reducing the resistance of the circuit and the mass bit. Further, an optimum discharge time was found when the capacitance and the inductance were varied. A low initial inductance increases the thrust efficiency. Experimentally, these trends can be verified by comparing two thrusters: SIMP-LEX and ADD SIMP-LEX, with their different initial inductances. For ADD SIMP-LEX, the optimal thrust efficiency for different capacities was determined to be 31% at 60 μ F for a 17 J configuration.

Review of Pulsed Plasma Thruster Development at IRS

The research and development process of the ablative pulsed plasma thruster SIMP-LEX is summarized. By experiment, the magnetic field in between the electrodes was measured with an induction probe, and a high-speed camera was applied to estimate the propagation of the plasma plume. Further, the optimization process towards higher thrust efficiency by means of geometric and electrical modifications is described. By analysis, the calculation of magnetic field and change in inductance is summarized and the numerical models for the performance simulation of SIMP-LEX introduced. A peak value of 0.7 T was found for the magnetic field created during an 80 J discharge whereas a propagation velocity of 28 km/s was estimated for the first plasma plume. The optimization process towards a new engineering model yielded an increase of 40% in mean exhaust velocity to 26 km/s, and a maximum thrust efficiency of 32% was evaluated respectively. Calculation of magnetic field and change in inductance based on the Biot-Savart law yield input for the slug model, which can then predict the general performance parameters with a good accordance.

Investigation of the Magnetic Field in a Pulsed Plasma Thruster

The magnetic field between the electrodes of an 80 J ablative pulsed plasma thruster was both measured and determined analytically. This was done to better understand the acceleration process and assess the accuracy of the analytical method. The measurements at different positions in both breech-fed and side-fed propellant configurations were performed using a shielded induction probe of 1 mm in diameter. To calibrate the probe, a Helmholtz coil was built. The magnetic field was calculated using the law of Biot-Savart and assuming a current sheet thickness of 3 mm. The measured magnetic field showed an overall peak at 0.7 T. It was possible to confirm the induced current loop formed between the electrodes at the time when the overall current passes through zero. The comparison between the magnetic field model and the measurements at the propellant surface of a breech-fed thruster showed reasonably good accordance.

Measurement of electron density by Stark broadening in an ablative pulsed plasma thruster

Electron density was measured by Stark broadening in an ablative pulsed plasma thruster. The asymmetrical deconvolution is used to obtain Stark broadening. The result shows that the electron density in the discharge channel is 2.534×1022 m−3 when the discharge energy is 5 J and the measured electron temperature is 18 000 K, and it is in excellent agreement with other experimental and theoretical data. The electron density in the discharge channel increases very minimally with increasing discharge energy.

Plasma acceleration processes in an ablative pulsed plasma thruster

Plasma acceleration processes in an ablative pulsed plasma thruster (APPT) were investigated. APPTs are space propulsion options suitable for microspacecraft, and have recently attracted much attention because of their low electric power requirements and simple, compact propellant system. The plasma acceleration mechanism, however, has not been well understood. In the present work, emission spectroscopy, high speed photography, and magnetic field measurements are conducted inside the electrode channel of an APPT with rectangular geometry. The successive images of neutral particles and ions give us a comprehensive understanding of their behavior under electromagnetic acceleration. The magnetic field profile clarifies the location where the electromagnetic force takes effect. As a result, it is shown that high density, ablated neutral gas stays near the propellant surface, and only a fraction of the neutrals is converted into plasma and electromagnetically accelerated, leaving the residual neutrals behind.

Ablative PPT. New quality, new perspectives

An adequate physical model for the ablative pulsed plasma thrusters (APPT) processes was developed as a result of complex theoretical and experimental studies conducted by RIAME for the processes of plasma acceleration in the APPT and determination of plasma parameters, electric currents, magnetic fields, volumetric electromagnetic forces inside and outside the discharge channel. Possibility for reaching higher thrust efficiency values for such thrusters due to the transfer from oscillative discharge to a quasi-aperiodic discharge was proved by test. Such transfer favors better matching for the distributions of volumetric electromagnetic forces and accelerated plasma (propellant) in the thruster. Analysis for the ways to increase the APPT operating process efficiency is presented. An important role of the elements of external power supply circuit for increasing thruster efficiency is shown. The studies conducted allowed development for the models of highly efficient APPT of new generation characterized by specific thrust pulse of up to 15– 22 km / s and thrust efficiency of up to 25–35%, which differ by high stability of operating process, design, technologic and operational simplicity, and low cost. Development of flight models of a highly efficient APPT for the Russian small spacecraft of “Vulkan” type was started.

Ablative Z-Pinch Pulsed Plasma Thruster

The design, performance, and basic features of ablative pulsed plasma thrusters based on the z-pinch configuration are discussed through a series of experiments and numerical simulations. The motivation stems from the promise of the z-pinch configuration for increasing the thrust-to-power ratio and mass utilization efficiency above those of ablative thrusters with a conventional rectangular geometry. The performance of a series of ablative z-pinch pulsed plasma thrusters is characterized using a swinging-gate thrust stand and mass ablation measurements. The performance measurements are complemented by additional experimental diagnostics (current monitoring and high-speed photography) and numerical modeling in order to gain an understanding of the acceleration mechanism and provide direction for future design iterations. Three iterations in the design of the thruster result in thrust-topower ratios ranging from 12‐45 µN/W, with specific impulse and thrust efficiency values spanning 240‐760 s and 2‐9%, respectively. Numerical simulations show reasonable quantitative agreement with the experimental data and predict the existence of an optimal thrust chamber aspect ratio, which maximizes the thrust-to-power ratio.

Zero-Dimensional Model for Preliminary Design of Ablative Pulsed Plasma Teflon Thrusters

A new procedure to estimate the ablated mass in ablative pulsed plasma thrusters (APPTs) is presented throughout this paper. The procedure is based on considering the energy balance as related to the generation and transport of the current pulse, arc formation, heat dissipation by joule effect, and the plasma acceleration. This new balance equation is written in terms of the ablated mass of Teflon® as a function of the main electric and geometric parameters. The introduction of an ideal arc thickness in the descriptive equations allows the buildup of a simplified model well suited for the prediction of performances and for preliminary design purposes. The resulting analytical model shows the ability to reproduce trends and behaviors having already been observed theoretically as well as experimentally by other authors. On the quantitative side, this new model allows for improved performance predictions as compared with semi-empirical models. Comparison of the theoretical results obtained by means of this model for the impulse bit, with the experimental figures corresponding to five APPTs, results in an average relative error around 20%.

Pulsed thrust measurements using laser interferometry

An optical interferometric proximeter system (IPS) for measuring thrust and impulse bit of pulsed electric thrusters was developed. Unlike existing thrust stands, the IPS-based thrust stand offers the advantage of a single system that can yield electromagnetic interference-free, high accuracy (&amp;lt;2% error) thrust measurements within a very wide range of impulses (100 μN s to above 10 N s) covering the impulse range of all known pulsed plasma thrusters. In addition to pulsed thrusters, the IPS is theoretically shown to be capable of measuring steady-state thrust values as low as 20 μN for microthrusters such as the field emission electric propulsion thruster. The IPS-based thrust stand relies on measuring the dynamic response of a swinging arm using a two-sensor laser interferometer with 10 nm position accuracy. The wide application of the thrust stand is demonstrated with thrust measurements of an ablative pulsed plasma thruster and a quasi-steady magnetoplasmadynamic thruster.

System Optimization of Ablative Pulsed Plasma Thrusters for Stationkeeping

We present a realistic and easy-to-apply model for the system optimization of ablative pulsed plasma propulsion for Stationkeeping missions. The model yields the characteristics of the propulsion module(s) that lead to minimum propulsion system mass for given mission specifications. We show that using an empirical relation for mass production, an expression relating the optimal (minimum) propulsion system mass to the thruster's most important scaling parameter E/l (the ratio of the discharge energy E to the impulse bit /) can be found and, remarkably, is independent of the state of capacitor and power conditioning technologies. This independence allows unfolding the relations between the module mass, E/I, and the mission requirements in a single plot that is applicable for a wide range of delta-v and payload masses. The use of the model to characterize the optimal design of a pulsed plasma thruster system is illustrated with the example of a 10-year north-south Stationkeeping of a medium size commercial satellite at geosynchronous orbit. We show that even with off-the-shelf capacitors, the use of such thrusters can result in propulsion mass fractions as low as 6.4%. A c d E F(Ei) f g /v

Electromagnetic emission experiences using electric propulsion systems

Introduction A electric propulsion systems become ready to integrate with spacecraft systems, the characterization of propulsion system radiated emissions is of significant interest. This paper briefly summarizes communications and dynamic electromagnetic experiences using electromagnetic, electrostatic, and electrothermal propulsion systems. Electromagnetic radiated emission results from ground tests and flight experiences are presented, with particular attention paid to the performance of spacecraft subsystems and payloads during thruster operations. The impacts to transmission of radio frequency signals through plasma plumes are also reviewed. Over the last 30 years, more than 60 spacecraft were flown using electric propulsion systems for drag makeup, stationkeeping, and The major flight-qualified thruster systems are ablative pulsed plasma thrusters (PPT), ion thrusters, magnetoplasmadynamic (MPD) thrusters, and resistojets. Average power dedicated to the propulsion system ranged from about 3 W for early pulsed plasma devices to about 1 kW for each (space electric rocket test) SERT-II ion (See Table 1.) Table 1 does not characterize all electric propulsion flights; only publications that refer to thrusters electromagnetic (EM) emissions or RF interactions with plumes were considered in this survey. Except for some resistojet systems, most electric propulsion systems require power processing equipment to tailor the battery or solar array power to that of the thruster. The thruster/power processor produces an electromagnetic environment that could potentially impact spacecraft systems such as communications, guidance, navigation and control, payloads, and experiments. The EM environments may have permanent and varying magnetic fields along with radio frequency and conducted electrical emissions. Prior to flight, electromagnetic interference (EMI) measurements must be made, and hardware must be immunized or the level of EM I reduced to satisfy the compatibility requirements. Previous papers have surveyed the particle and field interactions using ion thrusters and MPD arcjets. This paper will review radio frequency interference (RFI) component test specifications, results of spacecraft integration tests, and radiated emissions from flight systems using electromagnetic, ion, and electrothermal thrusters. Spacecraft schematics and ground test configurations will be discussed, general test procedures will be synopsized, and spacecraft integration test results will be reported. Most of the EM emission results were obtained from four flights using ablative pulsed plasma thrusters, six ion propulsion flights, and three flights using electromagnetic devices (see Table 1). Only published or publicly available information was used in compiling this report. Most of the electric propulsion work pertaining to EM emissions characterization was performed in the U.S., Japan, USSR, West Germany, and China.