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Experimental Evaluation of a Micro Liquid Pulsed Plasma Thruster Concept

Daniel Simon*, H. Bruce Land†, and Jerold Emhoff‡ The Johns Hopkins University Applied Physics Laboratory, Laurel, MD 20723

Researchers at The Johns Hopkins University Applied Physics Laboratory (APL) have developed a novel miniaturized pulsed plasma thruster (PPT) concept that utilizes a liquid propellant. Such devices could theoretically provide ultra-low dry mass, high specific impulse (Isp), flexible propellant options, and long-term reliability in a single package. Experimental devices have been designed and fabricated to help evaluate the concept. Typically, these devices have been powered by 0.5 - 1.0 µF capacitors at 200 - 700 V. Electrical discharge characteristics have been measured and peak currents have been estimated to be about 1.9 kA for a device powered by a 0.937 µF capacitor at 600 V. Impulse bit measurements have been facilitated by a custom-built thrust stand. The thrust stand was calibrated by dropping small nylon balls on the thruster and measuring the momentum transfer with a high speed camera. Impulse bits from a thruster powered by a 0.937 µF capacitor over a 200 - 700 V range were found to vary from 0.4 – 0.6 µN-s. Additional testing and analysis techniques are planned to quantify the propellant mass consumed per impulse bit. These and further experiments can be used to evaluate and improve micro-liquid PPT (MILIPULT) devices.

I. Introduction

otential applications for micropropulsion technologies include primary and attitude control for miniature satellites, fine pointing and positioning control for conventional satellites, and distributed manipulation of

extended space structures. Depending upon the specific mission requirements, each of these potential applications will require different propulsion system characteristics. Sound performance will likely be required for any application. All other things being equal, a technology offering a higher Isp and electrical efficiency will always be preferable. However, other system characteristics such as peak power consumption, minimum impulse bit, and overall dry mass will often determine a technology’s suitability for a given mission. The characteristics of the propellant itself may even have considerable mission consequences.

Pulsed plasma technology has a number of attractive attributes that can be exploited in a miniature propulsion system. The classic solid-fueled PPT concept is illustrated in Figure 1. PPT systems have the longest history of successful flight applications amongst all electric propulsion devices, which provides a thorough database of implementation details. The technology is scalable and inherently pulsed in nature, enabling very low power operation and miniscule impulse bit generation. Mechanical simplicity makes these devices straightforward to manufacture, storable indefinitely, and durable in flight. Modern PPTs have typically employed Teflon propellant; however, the technology is compatible with a broad array of substances that could provide tailored plume characteristics. With minimal electronic controls they are capable of thrust variation. Furthermore, PPTs are capable of operating with very high specific impulses exceeding 1000 sec.

* Aerospace Engineer, Research and Technology Development Center, AIAA Senior Member. † Electrical Engineer, Research and Technology Development Center, AIAA Member. ‡ Post-Doctoral Researcher, Research and Technology Development Center, AIAA Member.

P

42nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit9 - 12 July 2006, Sacramento, California

AIAA 2006-4330

Copyright © 2006 by the American Institute of Aeronautics and Astronautics, Inc. Under the copyright claimed herein, the U.S. Government has a royalty-free license to exercise all rights for Governmental purposes. All other rights are reserved by the copy-right owner.<NULL>

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Figure 1. Illustration of conventional solid propellant Pulsed Plasma Thruster (PPT) concept. Recently, a handful of programs have begun to explore the potential of miniaturized pulsed plasma thruster

designs. Miniature two-dimensional PPTs with fairly conventional designs were built for the University of Washington’s Dawgstar satellite. Electrical discharges in these devices are provided by 1.3 µF capacitors charged to 2.8 kV (containing approximately 5.1 J). Each discharge generates an impulse bit of 60-66 µN-s, which yields a thrust-to-power ratio of approximately 12 µN/W. The entire flight-qualified propulsion system – consisting of a number of juxtaposed thrusters – weighs slightly more than 4.0 kg.1 Even smaller PPTs with coaxial electrodes were fabricated at the Air Force Research Laboratory (AFRL). A 3-electrode version was driven by 2 µF capacitors holding 800-1000 V (0.64 – 1.0 J). This research effort largely focused upon reliable device operation and massive reductions in dry mass.2 These devices are slated to fly on the forthcoming Air Force Academy satellite, FalconSat-3.

These previous experiments with miniature ablative PPTs have demonstrated that their dry mass could be significantly curtailed (via operation at moderately low voltages), that modest fabrication techniques could be used, and that near term application was feasible. Awadallah3 also showed that the low energy and small scale of these devices should naturally minimize the electromagnetic radiation (and potential for interference) from their recurring discharges. However, the extremely small devices built under the AFRL program began to expose a number of limitations imposed by the ablating solid propellant mechanism. For example, these devices must operate at a staggeringly high ratio of energy to propellant surface area in order to prevent charring of the Teflon propellant surface.4 A review of eleven conventional PPTs by Burton and Turchi5 shows that this ratio typically ranges from 0.5 to 10 J/cm2. In contrast, the AFRL micro-PPT operates at approximately 100 J/cm2. This is certain to elevate the propellant surface temperature, which has been noted to encourage the late-term propellant ablation phenomenon.6 This phenomenon consumes propellant without contributing much to thrust production and can significantly reduce the overall Isp. Moreover, in order to generate a reasonable amount of total impulse these devices must utilize long propellant sticks that result in an awkward system form factor.

II. Micro Liquid PPT Description

The full potential of miniature PPT technology has clearly not yet been demonstrated. The inherent

inefficiencies and operating constraints experienced by the coaxial solid-propellant micro-PPT indicate that alternative propellant management systems should be explored. A liquid-based micro-PPT would eliminate the solid propellant surface that is prone to charring. This would allow the microthruster to operate reliably over a wider range of energy levels (specifically lower energy levels) and enable lower power operation. Other benefits of a liquid-based micro-PPT system include: constant thruster geometry, conveniently shaped propellant storage vessels, propellant sharing between thrusters, controllable power-to-mass bit operation, and propellant flexibility.

To investigate the benefits of liquid-based micro-PPTs, researchers at APL have developed experimental miniature devices that utilize water for propellant. The use of water for propellant in a PPT is not without precedent. Scharlemann has developed a conventional-scale (2.54 cm electrode gap) PPT that could operate using either solid Teflon or vaporized water. The thruster operated at energy levels ranging from 10-30 J per discharge. Experiments showed that the thruster performance was consistently better with the water propellant than with Teflon.7 Ziemer developed several conventional-scale PPTs that could operate using argon or vaporized water. The thrusters performed a succession of high-frequency discharge bursts (to completely utilize each injected gas bit). Each discharge typically contained 2-4 J of energy. Experiments showed that the PPTs generally performed better with the water propellant.8

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Experimental Micro Liquid Pulsed Plasma Thrusters (MILIPULT) devices have been fabricated at APL from a lightweight fiberglass substrate using printed circuit board (PCB) techniques. These processes afford tight tolerances, offer excellent repeatability, and could enable eventual mass production. The experimental units currently being examined are fabricated in batches of 15. Deliberate design variations are typically distributed within these batches. A photo of a first-generation micro liquid-PPT device is shown in Figure 2. It measures approximately 25mm x 25mm x 13mm and weighs about 13.5g (without its power electronics). The electrodes (approximately 1 mm wide by 3.2 mm long) are arranged in a planar (2-D) fashion and spaced 0.8 mm apart. A more detailed description of their internal construction and fabrication approach has been provided in earlier publications9

Figure 2. Typical Micro Liquid-PPT (MILIPULT) experimental device.

The MILIPULT devices have a small reservoir of liquid water within their structure. The first generation

of devices has minimal propellant control and provides a steady flow of water vapor through the device. Several flow path variations have been explored, which provide a range of constant flow rates from device to device. These variations allow the experimental evaluation of device performance with different mass bits participating in the impulses. These prototypes are useful for examining basic device behavior; however, they are not representative of the ultimate capabilities being sought. Most importantly, they do not efficiently utilize the propellant passing through them. Additional device features are currently being investigated which would significantly reduce propellant flow between discharges and provide variable propellant flow rates in a single device.

Figure 3. MILIPULT experimental sustain and trigger circuits.

The MILIPULT devices are operated via fairly traditional PPT power processing and ignition circuits, as

illustrated in Figure 3. A sustain capacitor connected to the device’s two parallel electrodes is charged to a

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moderately high voltage through a high voltage power supply. Then a high voltage trigger pulse is sent to a spark gap embedded in the device by dumping a secondary capacitor across a step-up transformer. To minimize inductance in the main discharge circuit, the sustain capacitor is mounted directly on the device surface. To simplify laboratory testing, the remaining electronics are typically enclosed in modular project boxes operating remotely (outside the vacuum chamber). To explore the device’s behavior a variety of sustain capacitors have been employed in testing. Three promising capacitors used in laboratory testing include the Panasonic 630V/0.56µF polyester leaded capacitor (Part# ECQ-E6564KF), TDK 630V/0.47µF multi-layer ceramic capacitor (Part# CKG57NX7R2J474M), and Vishay 500V/0.68µF multi-layer ceramic capacitor (Part# VJ3640Y684KXEAT).

III. Micro Liquid PPT Operating Characteristics

The average propellant mass flow through these devices has been estimated by measuring their propellant

mass utilization over a fixed time period. Since these experimental devices do not attempt to control propellant flow between discharges, their average propellant mass flow rates are higher than necessary for their relatively slow repetition rates (typically 1 Hz). Currently, MILIPULT devices with two different mass flows are being investigated: approximately 10 and 20 µg/s.

The electrical discharge behavior of these devices is currently monitored by measuring the voltage drop across the sustain capacitor with a (high impedance) high voltage probe. Voltage traces of the sustain discharge are recorded as a function of time and fitted to a simple “lumped parameter” LCR circuit model. The plot in Figure 4 shows a typical discharge for a 20 µg/s unit powered by two TDK 0.47 µF capacitors connected in parallel and charged to 600 V. The discharge shape appears to be just slightly underdamped. The circuit capacitance was measured to be 0.937 µF using a digital multimeter. The fitted circuit resistance and inductance were 290 mΩ and 43 nH, respectively. The LCR model predicts the peak current to be about 1.9 kA; however, it is noted that the model initially lags the actual discharge (and likely underestimates the peak current).

-200-100

0100200300400500600700800

0.0E+00 5.0E-07 1.0E-06 1.5E-06Time (sec)

Cap

acito

r Vol

tage

(V)

ExperimentModel

Figure 4. Typical sustain capacitor discharge as a function of time

and a LCR circuit model fitted to the data.

The same MILIPULT device was operated over a range of initial sustain voltages from 200 to 700 V (slightly above the capacitors’ rated voltage limit). Repeated pulses at a single operating condition were observed to produce extremely repeatable discharge traces. LCR models were fitted to each discharge and summarized in Table 1. Although this simple circuit model may not provide a perfect quantitative measure of the thruster circuit parameters, it does clearly indicate trends in behavior. As the starting voltage on the capacitor is increased, the total circuit resistance appears to decrease from about 370 mΩ to 270 mΩ, the total circuit inductance appears to decrease from about 110 nH to 41 nH, and the peak current increases from about 360 A to 2.3 kA.

The resistance in the thruster discharge circuit has contributions from the plasma layer, the printed copper conductors, and the capacitor. The drop in resistance at higher voltage levels is likely due to hotter, more conductive plasma sheets generated by the increased current levels. Measurements of the equivalent series resistance across the capacitor were made at various frequencies with an Agilent 4395A Network/Spectrum/Impedance Analyzer. These measurements indicated that the two capacitors contribute little (about 8 mΩ total in their parallel configuration) to the overall circuit resistance. Alternatively, calculations of the resistance in the thruster’s printed copper traces indicate that they could contribute almost 70% of the total circuit

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resistance predicted for the higher voltage operating conditions. Modified MILIPULT device designs with enlarged copper traces are underway that should significantly reduce the circuit resistance and increase the current flow.

Table 1. LCR circuit model parameters for a single device

operating over a range of sustain capacitor voltages

Vinit (V) R (ohm) L (nH) C (µF) Ipeak (kA) 700 0.27 41 0.937 2.3 600 0.30 46 0.937 1.9 500 0.32 49 0.937 1.5 400 0.33 55 0.937 1.2 300 0.35 70 0.937 0.73 200 0.37 110 0.937 0.36

The apparent increase in the circuit inductance is unexpected. Estimates of the inductance in the copper

traces account for nearly all the inductance predicted by the lumped LCR model at the higher operating voltages (approx. 40 nH). The circuit inductance is fixed by the circuit geometry and should not vary with operating conditions. Continued analysis is needed to determine if this result is an artifact of this simple modeling process or an authentic representation of the device behavior.

IV. Micro-Liquid PPT Impulse Bit Measurements

Thrust Stand Design In order to understand and improve the capability of the MILIPULT devices, it is crucial to accurately

measure their delivered thrust. Since the MILIPULT devices are pulsed in nature, they actually deliver a series of minuscule (approximately 1 µN-s) impulse bits. A miniature thrust stand system was designed at APL that is capable of resolving each of these individual impulses. This allows the average thrust production and ultimately the shot-to-shot variation to be determined for any set of thruster operating conditions.

Figure 5. MILIPULT device mounted on an

APL thrust stand flexure.

Figure 6. Assembled thrust stand system with

MILIPULT device and ball dropper calibration tool

As shown in Figure 5, the thrust stand system uses a stiff flexure comprised of a 3 x 3 cm platform suspended by four 6.35 x 8.30 mm cantilever beams in the middle of a rectangular frame. This flexure is fabricated in one piece from a sheet of 0.254mm (0.010”) stainless steel shim stock. A MILIPULT device is mounted on the moveable platform with its exit perpendicular to the flexure surface. Three strands of fine magnet wire (with an insulating shellac coating) are used to run power from (electrically isolated) copper pads on the flexure’s outer frame to the thruster. One wire serves to charge the PPT main capacitor, another carries a trigger pulse to the thruster, and a third serves as a common ground for both the sustain and trigger circuits. As shown in Figure 6, the flexure’s outer frame is clamped securely in a Plexiglas® stand allowing the suspended platform and thruster to vibrate freely. Spring-loaded IDI contact probes embedded in the Plexiglas® stand convey power from feed-throughs in the vacuum chamber to the copper pads on the

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flexure’s outer frame. Two separate Plexiglas® stands have been constructed that allow the steel flexure to be held either vertically (allowing the thruster to fire sideways) or horizontally (allowing the thruster to fire upwards). The stands are mounted on custom-built shaker tables driven by piezoelectric stack actuators. These tables can be used to vibrate the steel flexures in a controlled fashion (e.g. to search for resonant frequencies). This entire thrust stand assembly is small enough to operate within a benchtop vacuum chamber measuring 18.4 x 15.2 x 13.3 cm (7.25 x 6 x 5.25 in).

Laser Diode

Stationary Mirror

MoveablePlatform

VacuumChamber

Detector

Thrust StandFootprint

WindowBeam Splitter

Floating Optical Bench

Figure 7. Laser interferometer arrangement for measuring thrust stand vibrations.

An impulse imparted from the micro-PPT devices causes the platform to vibrate ever so slightly.

To measure the amplitude of these vibrations a Michaelson interferometer setup is employed. As illustrated in Figure 7, the beam from a 4.5 mW, 670 nm laser diode module from ThorLabs is passed through a non-polarizing beam splitter. One portion of the beam then passes through a quartz window into the vacuum chamber, reflects off a mirror mounted on the thrust stand’s moveable platform, and returns to the beam splitter. The other portion of the beam is reflected back to the beam splitter through a stationary reference mirror. The beam splitter then directs both beams towards a photodiode detector where they generate an interference pattern. As the position of the thrust stand platform moves, the intensity of the light at the detector varies as

⎥⎦

⎤⎢⎣

⎡⎟⎠⎞

⎜⎝⎛+=

λπ )(4cos1

2)( max txItI (1)

where x(t) is the position of the mirror and λ is the wavelength of the laser source. Note that the output is sinusoidal and repeats every time the platform moves by ¼ of a wavelength (167.5 nm). To minimize background vibrations due to environmental effects, the vacuum chamber and interferometer have been assembled on an optical bench with pneumatic isolation legs.

When excited by a brief impulse, the thrust stand’s steel flexure is designed to vibrate with a decaying sinusoidal pattern characteristic of a classic damped spring-mass system:

( )tm

Ietx effdeffdeff

bitteffneff,

,sin)( , ω

ωωξ−= (2)

where Ibit is the impulse bit, meff is the effective system mass, ξeff is the effective damping coefficient, ωn,eff is the effective natural frequency, and ξωω −= 1,, effneffd is the damped effective natural frequency. Since the damping coefficient is relatively small for this system, the oscillations decay relatively slowly. As such, the size of an impulse bit can be deduced from the amplitude of oscillations it generates. Previously, the system had been calibrated using a shifting resonance technique to infer the mechanical characteristics of the vibrating flexure by assuming that it behaved like a well-behaved single-degree-of-freedom system10.

To provide more confidence in this impulse bit data, a more direct calibration method was desired. A pneumatically-actuated device, which is visible in Figure 6, was designed and constructed to drop small

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(~1 mm diameter) balls on the thruster surface from variable heights. A high speed camera is used to capture the ball’s impact and calculate its incoming and outgoing velocity. Knowledge of the ball’s mass allows us to calculate the momentum change imparted to the ball. We assume that an equal and opposite momentum change (i.e. impulse) is imparted to the thruster. Nylon (φ=1.20 mm) and Delrin (φ =1.60 mm) balls have been used. The average mass of these balls (1.05x10-3 g and 2.95x10-3 g, respectively) was determined by collectively weighing a group of 100 on an electronic balance. So far, impulse bits as small as 0.5 µΝ-sec have been generated using this dropper mechanism. By simultaneously measuring the amplitude of the thrust stand platform’s vibration response to a variety of impulse bits, we can perform an accurate thrust stand calibration. Thrust Stand Shakedown The initial data collected with this system indicated that some further modifications to our equipment and procedures were necessary. We had previously reported that the natural frequency of the thrust stand vibrations tended to vary each time a new experiment was assembled. Furthermore, the frequency was often lower than expected. Initial models of the system had assumed the vibrating platform was inflexible because its cross section is larger than the cantilevers and thruster devices were firmly bonded to it. Initially, vinyl acetate (thermoplastic) droplets were used to bond the thrusters to the platform. A finite element structural model of the cantilevers, platform, thermoplastic droplet, and thruster showed that substantial flexing of the platform was occurring. Moreover, the model showed that variations in the thermoplastic droplet size strongly affected the resonant frequency of the system. To provide a stiff system with consistent behavior, we experimented with modified flexures and 2 mm thick G10 (glass epoxy sheet) stiffening plates assembled with metal or plastic hardware. Plastic screws and nuts exhibited too much elasticity. Metal screws and nuts increased the vibrating platform’s mass and reduced the system’s sensitivity to thruster impulses. Moreover, the added mass reduced the system’s natural frequency and appeared to make it more susceptible to background vibrations. The plot in Figure 8 shows the thrust stand response to a MILIPULT discharge using a stiffener plate assembled with metal nuts and screws. This experiment captured a low frequency (173 Hz) response from the thruster discharge. However, the system also exhibits continuous high frequency (>900Hz) vibrations. At this point, a Varian V250 turbo pump operating at 56,000 rpm (933Hz) was mounted in close proximity to our vacuum chamber.

3

3.5

4

4.5

5

5.5

6

6.5

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Inte

rfer

omet

erO

utpu

t (V)

0

100

200

300

400

500

600

700

Sust

ain

Cap

acito

r (V)

InterferometerSustain Capacitor

Figure 8. Representative vibrations of thrust stand platform

and stiffener plate assembled with metal screws and nuts (using the turbo pump).

Although our background noise levels tended to fluctuate in amplitude and frequency, these experiments indicated that the turbo pump was a primary contributor. In a bid to improve the data quality from the thrust stand, we replaced the turbo pump with an Innotec R-220 diffusion pump. This arrangement provided very clean data and new insight into the mechanics of the thrust stand. One experiment was performed in which a thruster device was epoxy bonded directly to a flexure. Typical vibration patterns in response to thruster firings, such as the one in Figure 9, contained a clear “beating” pattern. Such a pattern can result from the superposition of two vibration modes with close (but slightly different) natural frequencies. The pattern in Figure 9 could be elicited from a system with vibration modes at 180 and 230 Hz.

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2

3

4

5

6

7

-0.05 0 0.05 0.1 0.15Time (sec)

Inte

rfer

omet

erO

utpu

t (V)

Figure 9. Representative vibrations of the platform with the thruster device

bonded directly (without balancing) on using 5-minute epoxy (using the diffusion pump). To aid our understanding of the thrust stand’s vibrating structures, a series of 3-D finite element

structural models were created using the FEMLAB software package. The first model (Model A) consisted solely of the steel platform and cantilevers (E = 210 GPa, ν = 0.33, ρ = 7900 kg/m3). A modal analysis using this model predicted its first natural frequency occurred at 520 Hz. A bare flexure was clamped in the thrust stand and vibrated using the shaker table. The first resonant mode was found to occur at 527 Hz, which corroborated the model results. Next, a model (Model B) was created consisting of the steel platform/cantilevers and a 3x3 cm G10 stiffening plate (E = 16 GPa, ν = 0.33, ρ = 3700 kg/m3). A 0.254mm (0.010”) adhesion layer was included in the model between the stiffening plate and steel platform. Experiments had shown that this system’s first natural frequency occurred around 320 Hz. In order to match this value with the FEMLAB model, the elastic modulus of the adhesion layer had to be set to 0.1 MPa.

Finally, a third model (Model C) was constructed with a thruster device bonded directly to the vibrating platform (the system that created the beating signal in Figure 9). The thruster device was modeled as a dimensionally-accurate block (25.4 x 25.4 x 20 mm) with a uniform density (1160 kg/m3) that provides the correct total mass for the device. The adhesion layer between the thruster and the platform was given the same properties as were used in Model B. A modal analysis showed that the first 3 vibration modes fell into a relatively small frequency range. The 2nd and 3rd modes, whose mode shapes are shown in Figure 10, arose at 190 Hz and 220 Hz. Although they don’t perfectly match the frequencies suspected of generating the beating pattern in Figure 9, they are a reasonably close prediction.

a) b) Figure 10. 2nd (a) and 3rd (b) mode shapes predicted by FEMLAB for a model of a

thruster device bonded to directly to a flexure with 5-minute epoxy.

Note that the 2nd and 3rd modes shapes illustrated in Figure 10 correspond to a translation and rotation of the thruster device. Also note that the thruster exit is not centered in our current experimental devices. In fact, all experiments performed up to this point had aligned the thruster so that its exit was skewed towards one set of cantilever arms. It is understandable how the device’s impulse would encourage the rotational motion embodied by the 3rd mode shape. If a combination of vibration modes were to blame

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for the irregular patterns sometimes produced by the thrust stand, the effect could be reduced by centering thruster exit on the platform and balancing the vibrating mass. MILIPULT Measurements To provide more regular vibration patterns from the thrust stand, a balanced flexure assembly was prepared. A 3 x 3 cm square of 2 mm thick G10 was epoxy-bonded to the platform. A MILIPULT device (the 20 µg/s version) was epoxy-bonded to the G10 square so that its exit was centered on the platform. The G10 square served to uniformly stiffen the platform and prevent the thruster from contacting the cantilever arms (since the thruster structure hung over the edge of the platform). This assembly was then balanced on a knife edge by adding ballast to the other side of the platform. This assembly was tested in the thrust stand using the low vibration diffusion pump. Although the diffusion pump had a nominal pump speed equivalent to our prior turbo pump (which could achieve 10-4 - 10-5 torr chamber pressures), it was only able to achieve chamber pressures of 10-3 torr. Facility improvements are underway to improve these conditions for future tests. However, the balanced flexure assembly did provide a reasonable improvement in the character of the thrust stand impulse responses. The plot in Figure 11 displays a typical vibration pattern elicited from a thruster impulse. The response is not exactly a perfect damped sinusoid, but the strong beating pattern has been eliminated.

-40

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-20

-10

0

10

20

30

40

-0.025 0 0.025 0.05Time (sec)

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st S

tand

Vib

ratio

n (n

m)

0

100

200

300

400

500

600

700

800

Sust

ain

Cap

acito

r (V)

InterferometerSustain Capacitor

Figure 11. Typical thrust stand vibration in response to a

thruster impulse using a balanced platform assembly.

A series of ball drops was performed to calibrate this system. To maintain a balanced system, the balls were dropped as close to the thruster exit as possible. Note that these vibration patterns exhibited imperfections similar to those observed in the thruster impulse data. Since the vibration patterns were not perfectly regular, it was uncertain how to “best” interpret the magnitude of each vibration. Consequently, two metrics in the vibration patterns were recorded for each ball drop: 1) the amplitude of the first peak and 2) the maximum peak-to-peak amplitude. Figure 12 shows the ball impulse bit plotted as a function of these two vibration amplitude metrics. A least squares linear curve fit was performed on each data set. The average absolute error between either of these curve fits and the measured impulse bits was the about the same: 0.045 µN-s. Additionally, in both cases the standard deviation of the absolute error values was 0.065 µN-s. This implies that an individual ball impulse bit could be predicted to within 0.13 µN-s with a 95% confidence level from either vibration amplitude metric.

A series of MILIPULT firings were then performed over a range of voltage levels from 200V to 700V on the device’s sustain capacitors. These tests were conducted with two 0.47µF capacitors connected in parallel (measured capacitance was .937 µF) providing a range of discharge energies from .018 to 0.22 J. Using the measured maximum peak-to-peak amplitude in the thrust stand’s response and the corresponding ball drop calibration, 5 impulse bit values were measured at 6 discrete discharge energy levels (30 measurements total).

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0.40.450.5

0.550.6

0.650.7

0.750.8

0.850.9

0 10 20 30 40Thrust Stand Response (nm)

Impu

lse

Bit

( µN

-s)

First PeakMax Peak-to-Peak

Figure 12. Impulse bits imparted by ball drops versus two metrics in the vibration response:

1) amplitude of the first peak and the 2) max peak-to-peak amplitude

y = 0.7098x0.1259

R2 = 0.9986

0.3

0.4

0.5

0.6

0.7

0 0.05 0.1 0.15 0.2 0.25Discharge Energy (J)

Impu

lse

Bit

( µN

-s)

Individual Shots

Average Values

Figure 13. Impulse bit measurements for a thruster powered by

937 µF of capacitance at 200 to 700 V. The plot in Figure 13 shows that the impulse bit varied between 0.4 and 0.6 µN-s. Given the

uncertainty in the current calibration curves, it is impossible to definitively evaluate the shot-to-shot variability. Nonetheless, the thrust data seems to exhibit less scatter at the lower energy levels than at the higher ones. The measured impulse bits clearly increases with energy; however, the relationship does not appear to be linear. Both logarithmic and power law functions fit well to the data. At the lowest energy level, the average impulse-to-energy ratio was 24.4 µN-s/J. At the highest energy level, the average value was only 2.6 µN-s/J. Physical mechanisms that could explain this behavior are being studied.

The data collected using the existing MILIPULT devices has provided useful insight into their operating behavior. Modified MILIPULT devices have been designed that center the thruster exit on the structure. This will greatly simplify the process of balancing the thrust stand platform and hopefully improve the regularity of the platform vibration patterns. Furthermore, the absence of ballast on the thrust stand should provide larger amplitude vibrations (stronger signals) in response to impulse excitations. These changes should help reduce any measurement uncertainty being induced by the thrust stand itself and may improve our shot-to-shot resolution. These devices were not available for testing prior to the submission of this paper, but forthcoming tests will be described in future publications.

V. Summary

A novel micro-PPT concept is being explored that consumes a liquid propellant. Previous

examinations of micro-PPTs using a stationary solid propellant have demonstrated that the technology is capable of producing compact and reliable micropropulsion devices. However, reliance upon previous solid propellant designs reduces electrical efficiency, increases minimum power requirements, and imposes an awkward thruster form factor. In theory, a liquid-based micro-PPT could retain the simplicity and ruggedness of the PPT concept while eliminating the constraints associated with the solid propellant.

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To investigate this claim, researchers at the APL have developed a number of laboratory prototype devices based upon the micro liquid-PPT concept. The devices weigh about 13.5 g and have 2-D electrodes spaced about 0.8 mm apart. Several flow path variations have been incorporated in these devices to provide a handful of propellant flow rates for testing. Current devices allow propellant to flow continuously (during and in between pulses); however, mechanisms that will provide pulsed propellant flow are under development. Two particular device designs with flow rates of approximately 10 and 20 µg/s are the subject of current study. Electrical discharge and impulse bit measurements have been made with devices powered by capacitances ranging from 0.5 to 1.0 µF and voltages ranging from 200 to 700 V.

A miniature thrust stand and ball dropping mechanism for calibration have been developed to measure impulse bits from these devices. For a device with a steady 20 µg/s propellant flow, the impulse bit was measured in response to discharges ranging in energy from .018 J to 0.22 J. The resulting impulse bit increased in a non-linear fashion from 2.6 to 24.4 µN-s/J.

Further testing and analysis are planned to determine the nature of this non-linear dependence and to quantify the propellant mass bit participating in each impulse. Eventually, improvements in the thruster specific impulse and efficiency can be pursued by optimally matching the electrical discharge and mass flow conditions. To accomplish these goals, more sophisticated hardware designs have been conceived that could provide real-time propellant flow control in an equally lightweight and low-power package.

References

1 Rayburn, C.D., Campbell, M.E., and Mattick A.T., “Pulsed Plasma Thruster System for

Microsatellites,” Journal of Spacecraft and Rockets, Vol. 42, No. 1, Jan-Feb 2005, p.161-170. 2 Gulczinski, F.S., et al., “Micropropulsion Research at AFRL”, 38th AIAA Joint Propulsion

Conference & Exhibit, Huntsville, AL, 16-19 July 2000. 3 Awadallah, R., et al., “Electromagnetic Emission Modeling for Micro Pulsed Plasma Thrusters,” 41st

AIAA Joint Propulsion Conference, Tucson, AZ, 10-13 July 2005. 4 Keidar, M., et al., “Propellant Charring in Pulsed Plasma Thrusters,” Journal of Propulsion and

Power, Vol. 20, No. 6, Nov-Dec 2004, p. 978-984. 5 Burton, R.L. and Turchi, P.J., “Pulsed Plasma Thruster,” Journal of Propulsion and Power, Vol. 14,

No. 5, 1998. 6 Spanjers, G.G., et al., “The Effect of Propellant Temperature on Efficiency in the Pulsed Plasma

Thruster,” Journal of Propulsion and Power, Vol. 14, No. 5, 1997. 7 Scharlemann, C.A. and York T.M., “Pulsed Plasma Thruster Using Water Propellant, Part II:

Thruster Operation and Performance Evaluation,” 39th AIAA Joint Propulsion Conference & Exhibit, Huntsville, AL, 20-23 July 2003.

8 Ziemer, J.K. and Petr R.A., “Performance of Gas Fed Pulsed Plasma Thrusters Using Water Vapor Propellant,” 38th AIAA Joint Propulsion Conference & Exhibit, Indianapolis, IN, 7-10 July 2002.

9 Simon, D.H. and Land, H.B., “Micro Pulsed Plasma Thruster Technology Development,” 40th AIAA Joint Propulsion Conference, Fort Lauderdale, FL, 11-14 July 2004.

10 Simon, D.H. and Land, H.B., “Instrumentation Development for Micro Pulsed Plasma Thruster Experiments,” 41st AIAA Joint Propulsion Conference, Tucson, AZ, 10-13 July 2005.