Pulsed Plasma Propulsion for a Small Satellite: Mission COMPASS P3OINT

Home , Cold gas thruster, Pulsed plasma thruster

Pulsed Plasma Propulsion for a Small Satellite: Mission 3 COMPASS P OINT∗

J.K. Ziemer†, E.A. Cubbin‡and E.Y. Choueiri§ Electric Propulsion and Plasma Dynamics Laboratory (EPPDyL) MAE Dept. Princeton University Princeton, New Jersey 08544

V. Oraevsky¶and V. Dokukink IZMIRAN, Moscow, Russia

AIAA-96-3292∗∗

Abstract Operational modes include a combination of right and left nozzle pulse sequences for determining the per- Pre-launch characterization and preparation of a Lin- formance and impulse bit of the APPT module in coln Experimental Satellite (LES 8/9) ablative pulsed space. Documented pre-flight experiments include plasma thruster (APPT) module for the COMPASS a temporary hardware addition that simulates the 3 3 P OINT Mission are described. COMPASS P OINT plasma discharge necessary for the pre-flight atmo- is a joint project between the Electric Propulsion and spheric integration test at IZMIRAN. Also described Plasma Dynamics Lab (EPPDyL) of Princeton Uni- are the experimental measurements of the thrust effi- versity and the Institute of Terrestrial Magnetism, ciency, center of mass, and moments of inertia of the Ionosphere and Radio Wave Propagation of the Rus- module. sian Academy of Science (IZMIRAN) to carry in- orbit investigations with an ablative pulsed plasma thruster onboard COMPASS, an IZMIRAN scientific 1 Introduction microsatellite to be launched in October, 1996. The unmodified LES 8/9 APPT module produces impulse Over the past four years there has been a resurgence bits of 285 µN-s at an Isp of 836 s using 25 W of of interest[1] in ablative pulsed plasma thrusters power on average. The APPT module will be used (APPT’s) due to a trend towards smaller satellites in to conduct in-orbit investigations of pulsed plasma the military, commercial and scientific sectors. Quite propulsion as well as to provide attitude control of often these small satellites are power-limited to below the satellite and a source of plasma for active space 300 W. To date, only APPT’s have the demonstrated experiments. Details of the power, command, and ability of providing the mass-savings advantages of telemetry signals required to operate the device on high specific impulse electric propulsion at and be- the COMPASS satellite are presented. Thermal con- low these power levels. trol and various operational modes are also outlined. Research and development efforts on APPT peaked in the mid-seventies but hopes for the availability of ∗Support from the Air Force Office of Scientific Research, grant number: F49620-95-1-0291 high power in space caused such efforts to all but †Graduate Student, Research Assistant. cease in the eighties. Recent studies have demon- ‡Graduate Student, Research Assistant. strated that even with off-the-shelf power storage and §Chief Investigator at EPPDyL. Assistant Professor, Applied power processing technologies, the APPT can offer Physics Group. Senior Member AIAA. substantial mass-savings for common stationkeeping ¶Director, IZMIRAN. kChief Investigator, IZMIRAN. missions[2]. nd ∗∗Presented at the 32 AIAA Joint Propulsion Conference The state-of-the-art of flight-ready APPT technol-

1 2 ZIEMER, CUBBIN, CHOUEIRI, ORAEVSKY & DOKUKIN: COMPASS P3OINT MISSION ogy of the mid-seventies is embodied in the Lin- 2 COMPASS and the COM- coln Experimental Satellites 8 and 9 Ablative Pulsed 3 Plasma Thruster (LES 8/9 APPT) modules[3] which PASS P OINT Mission were built but were never flown due to scheduling 2.1 Launch constraints. A great deal of research work went into the design and construction of these modules[4, 5] A 70 kg scientific microsatellite called COMPASS including previous spaceflights[6, 7]. Although the belonging to the Institute of Terrestrial Magnetism, LES 8/9 APPT modules did not go into space as Ionosphere and Radio Wave Propagation of the planned, a few Teflon APPTs have flown in space in Russian Academy of Science (IZMIRAN) will be the eighties[8]. launched later this year from a Russian Navy sub- As part of an AFOSR-sponsored program to ad- marine in the Barents sea using a 3-stage converted vance APPT technology to meet modern propulsion missile called Shtyl-2. The launch is scheduled for needs, Princeton University’s EPPDyL is carrying re- October 28, 1996 in connection with the celebration search on two complementary fronts: 1) Fundamen- of the 300th anniversary of the formation of the Rus- tal studies (theoretical and experimental) of APPT sian Navy. According to ref. [13], this will mark the ablation process, energy partitioning, scaling, circuit first time a converted military ballistic missile is used optimization and thrust efficiency, and 2) In-orbit in- to place a satellite in lower earth orbit (LEO). The vestigations of APPT performance. In the present satellite will be placed in a 400 km circular orbit with paper we discuss our efforts on the second front, par- an inclination of 78-79◦. The primary scientific goal ticularly the use of a LES 8/9 APPT module onboard of COMPASS is to conduct plasma studies related to a Russian scientific satellite called COMPASS to be earthquake precursors in the ionosphere. The satel- launched on October 28, 1996. lite, however, will also include a LES 8/9 APPT mod- The part of the COMPASS project that deals with ule supplied, conditioned and adapted to COMPASS the APPT is called COMPASS P3OINT for Pulsed by Princeton University’s EPPDyL. Plasma Propulsion in-Orbit Investigation and Navi- gation Test. In this paper we first describe the ob- 2.2 Scientific Goals of COMPASS jectives of the COMPASS P3OINT mission and gen- During its minimum lifetime of six months, COM- eral aspects of the COMPASS satellite. We then go PASS aims at the following scientific goals: into the details of the preparation and characterization of the APPT module. In particular, we define Studies of electromagnetic and plasma distur- the power, command, and telemetry signals necessary • bances in the ionosphere as precursors to earth- for the mission as well aspects of the required ther- quake, volcanic eruptions and other large-scale mal control for the thruster. The operational modes natural disasters. and pulsing sequences that will be used during the mission are also presented. Development of methods for the detection and • Before delivering the flight module to IZMIRAN monitoring of large-scale technological accidents for satellite integration in August of this year, three from space using microsatellites. instrumental steps have been accomplished. First, Investigation of the processes of electrodynamic temporary internal and external hardware modifica- • interaction between the atmosphere, ionosphere tions have been made to allow a simulated discharge and the Earth magnetosphere through passive at atmospheric pressure for a pre-flight satellite inte- and active space experiments. gration test. Second, the efficiency of the LES 8/9 APPT has been measured over 436 pulses yielding an average impulse bit of 285 µN-s within 5 µN-s, 2.3 General Objective of COMPASS 3 an efficiency of 6.8%, and a specific impulse of 836 s. P OINT These performance measurements were made using The part of the COMPASS project that deals with the interferometric proximeter system (IPS) as de- the APPT is called COMPASS P3OINT and has the scribed in ref [9, 10]. Third, all moments of inertia following objectives: have been found experimentally to within 1.5% using a bifilar technique described here in detail as well Provide attitude adjustment and stabilization as in ref [11, 12]. All three of these experiments are • for COMPASS (pitch angle control, instrument outlined in this paper. orientation, drag compensation, etc. ) ZIEMER, CUBBIN, CHOUEIRI, ORAEVSKY & DOKUKIN: COMPASS P3OINT MISSION 3

Provide a source of high velocity plasma for iono- Plasma • Diagnostics APPT Aerodynamic spheric perturbation and active space experi- Stabilizers ments on COMPASS.

Validation of APPT thrust and efficiency perfor- • mance baselines inferred from attitude changes caused by APPT firings. Equipment Thruster Validation of APPT operation baselines. Module Module • Investigation of on-orbit electromagnetic inter- • ference (EMI) from the APPT using the wave Figure 1: COMPASS Satellite in launch configura- diagnostics on COMPASS. tion. Provide on-orbit propulsion data to guide the • scientific part of EPPDyL’s AFOSR-sponsored Commands to COMPASS will be sent via the com- research on APPT’s. mand line starting at IZMIRAN through a reserved line from the city of Sochi. Macintosh computers at Provide hands-on astronautics and satellite in- IZMIRAN and Princeton linked through the Internet, • tegration education through the involvement of provide means for communication and digital data graduate and undergraduate students transfers. There are two modes of data collection: 1) Moni- 2.4 COMPASS Scientific Payload tor Mode and 2) Data Recording Mode into the onboard 32 MB memory. The data is transmitted via COMPASS has a suite of diagnostics intended for its the telemetry channel to the ground receiver stations scientific investigations. These instruments will be at IZMIRAN in the city of Troitsk in the Moscow operating during the APPT firings providing scien- region. Another telemetry link will be to the city of tific data about wave emission and particle energetics Sochi. The data transmission rate exceeds 1 Mbit/s during the COMPASS P3OINT mission. The instru- and the carrier frequency is 1.7 GHZ. ments include:

High frequency (HF) wave system for measuring 2.6 The Structure of COMPASS • waves in the range of 0.1-15 MHz COMPASS consists of two main modules: an An ultralow frequency (ULF)/ extremely low fre- equipment module and a thruster module. Before • quency (ELF) wave system for measuring in the launch, the two modules are collapsed into one over- 0.1 Hz-23 kHz range. all structure that fits inside a conical canister with an elliptical bottom section. This configuration is neces- Plasma diagnostics package for measuring sitated by the limited space available in place of the • plasma parameters warhead of the Shtyl-2 converted missile. Optical system consisting of two photometers Schematics of COMPASS before and after deploy- • and a TV camera ment are shown in figures 1 and 2 respectively. Af- ter the capsule separates from the launcher, the de- A 3-axis DC magnetometer ployment mechanism flips the thruster module 180 • degrees (see figure 2) and deploys the aerodynamic A particle spectrometer for high energy particles. stabilizers which are also covered on both sides with • solar arrays. 2.5 Satellite Power, Command, and The equipment module consists of two modular Telemetry panels: 1) a panel with scientific equipment and 2) a panel of heat sink radiators for the service equip- The power supply consists of a set of solar panels ment. In addition to these two panels, the equipment and a buffer accumulator battery with a capacity of module contains a longitudinal mechanical interface about 10 Ahr. During regular operation, the available through which the module is attached to the launch power will be 40 W at 27 V on average. capsule. 4 ZIEMER, CUBBIN, CHOUEIRI, ORAEVSKY & DOKUKIN: COMPASS P3OINT MISSION

3-D Magnetometer

Plasma Wave Diagnostics Aerodynamic UV Optical Stabilizers Photometer APPT Teflon Bars APPT Capacitor APPT Nozzles

Equipment Thruster Module Module

Electric Field Probe

Probe Assembly

1.4 m 2.3 m

Figure 2: COMPASS Satellite in deployed conﬁguration.

The thruster module consists of a stationary tion radiators expected to radiate 28 W into space frame with a movable gimbal assembly on which while maintaining the equipment panel at 293◦ K the thrusters and the aerodynamic stabilizers are within a range of 10◦ K under the least favorable mounted along with the mechanism that allows de- sunlight/darkness conditions. The heat rejection sys- ploying the stabilizers. The thrusters consist of one tem consists of an elliptical radiator with an area of cold gas thruster operating with xenon and the APPT .092 m2, a set of thermal accumulators using .174 kg module of COMPASS P3OINT described in detail in of ethyl alcohol and 4 heat pipes for enhanced heat section 3 below. conduction. The thermal management of the APPT A series of probes (Langmuir probe, magnetic and module is described in section 3.3. electric ﬁeld antennas), are mounted on retractable booms that are retracted before launch into their 2.8 Stabilization and Attitude Con- housing on both the equipment and thruster mod- trol Systems ules. The structure is made of aluminum alloy AMG-6 COMPASS will be placed initially in a 400 km orbit. which does not require any protective anti-corrosion The attitude control system is designed such that the coating. The modular panels with heat pipes are also microsatellite takes at least 6 months to degrade to made of aluminum alloys. The booms that carry the an altitude of 250 km. All systems onboard have a probes are made of plastics. The aerodynamic sta- lifetime exceeding 6 months. bilizers are made of honeycomb aluminum structure. A combination of gravitational and aerodynamic All movable parts have anti-corrosion coatings. stabilization will be used to maintain spacecraft ori- The entire satellite is subjected to outgassing with entation within 10 degrees. The attitude of the space- a temperature above 70◦ C for about three days. craft can be determined from the onboard solar sen- sors to within 1 degree. 2.7 Thermal Management Partial compensation for orbit degradation and further attitude control is supplied by a cold gas thruster The thermal management system for COMPASS ex- operating with xenon at a thrust level between .1 and cluding the APPT module is passive with heat rejec- 1 N and the APPT module which provides impulses of ZIEMER, CUBBIN, CHOUEIRI, ORAEVSKY & DOKUKIN: COMPASS P3OINT MISSION 5

285 µN-s at an Isp of 836 s using 25 W of power with capacitor to discharge across the electrode gap. The a duty cycle of 0.89 Hz. Description of the APPT completed plasma circuit ablates more Teflon and in- module, its preparation and pre-launch characteriza- duces a magnetic field. The field interacts with the tion are the subject of the rest of this paper. current sheet forcing it out of the thrust chamber producing the impulse bit. 2.9 Organizational Participation This section of the paper documents the important module characteristics necessary for spacecraft inte- While COMPASS P3OINT is a joint collaboration gration and space flight. The center of mass and the between IZMIRAN and Princeton University’s EP- moments of inertia for the thruster have been found PDyL, COMPASS is an IZMIRAN project involving experimentally and results are presented here. The the participation of scientists from the Former Soviet electrical components are described further including Union, USA, Bulgaria, Poland, Slovakia and Hun- the main capacitor circuit and the discharge initia- gary. Scientific payload integration is performed by tion circuit. Thermal management is also discussed IZMIRAN and M.V. Frunze Design Bureau Arsenal. in this section of the paper. The launcher and launch are the responsibilities of the State Rocket Center of the Makeev Design Bu- reau. 3.1 Center of Mass and Moment of Inertia Calculations 3 Description of APPT Mod- The center of mass location along with the defined x, y, and z-axis of the APPT are shown in the module ule drawings at the end of this paper, figures 10 and 11. A bifilar technique (described in more detail later in Twelve LES 8/9 APPT modules were originally built the appendix) was used to find the moments of iner- for station keeping and orbital maneuvering duties tia of the LES 8/9 APPT. The bifilar method con- aboard two Lincoln Lab Communications satellites sists of hanging an object by two long wires, twisting [3]. Drawings of the top, bottom, side, front, and the object a small amount, and measuring the pe- back of the APPT flight module can be found at the riod of oscillation about the vertical axis between the end of this paper in figures 7, 8, 9, 10, and 11. wires. Care must be taken to make sure the center The LES 8/9 APPT is parallel plate self-field of mass is exactly half way between the two wires plasma accelerator using Teflon, stored in a solid so that the rotation is about a constant vertical axis form, for propellant. It has two sets of electrodes, that includes the center of mass. To determine the each offset at a 30 degree angle from center. As shown full moment of inertia matrix, including products of in figures 7 and 8, the Teflon rods are held in place on inertia, six measurements on six different axis are re- the rear edge of the electrodes by a negator spring and quired. A coordinate transformation then produces retaining shoulder. Energy is stored in one 1500 V, the full matrix about the desired axis. For this analy- 17 µF silicon-mylar capacitor for a pulsed discharge sis, in cases where non-linear simultaneous equations that lasts 12 µs. Instantaneous power levels are on needed to be solved, they were solved numerically. the order of 10 MW with peak current values near The moment of inertia matrix, Io is, in kg m2, 20 kA. The average power consumption is close to 25 W at a pulse frequency near 1 Hz. The average 3.78e 2 5.45e 4 5.78e 3 impulse bit over the 107 pulse lifetime of the thruster − − − Io = 5.45e 4 5.63e 2 5.89e 3 (1) is rated at about 300 µN-s with a specific impulse  7.78e − 3 5.89e − 3 4.90e − 2  of 1000 s. Thrust efficiency is rated at about 7%[3]. − − − Actual performance of the COMPASS P3OINT flight   module has been determined and is presented later 3.1.1 Moment of Inertia Error Calculations in this paper. The pulse begins when the main capacitor is fully The errors on the above results were obtained by nu- charged and a TTL command signal is sent to the merically monitoring the solutions as the input val- discharge initiation (DI) circuitry. One of four spark ues were varied throughout the respective uncertainty plugs (each set of electrodes have two spark plugs, ranges. The percent uncertainty on all moments of one on the top and bottom of the cathode) fire, ab- inertia (the diagonal elements) was found to be 1.5%. lating a small amount of Teflon and allowing the main The uncertainty of the products of inertia is 5%. 6 ZIEMER, CUBBIN, CHOUEIRI, ORAEVSKY & DOKUKIN: COMPASS P3OINT MISSION

3.1.2 Center of Mass Misalignment +5V +625V The center of mass was located by hanging the thruster by a wire and noting that the center of mass PLUG must align with the wire. Two such experiments lo- cate the center of mass. Using the parallel axis theo- 2µF 0.02µF rem, the center of mass location must be known to an error of ² such that the biﬁlar pendulum is rotating about an axis a distance ² away from the x axis as described in section A.2. In this case the parallel axis theorem says that the error for a rigid body mass (m) 2 2 e T is m² . For our case, m² was much smaller then Ixx as determined in the experiment.

3.2 Electrical Components Figure 3: Fairchild Discharge Initiation (DI) circuit. The power processing unit used for charging the capacitors and the discharge initiation circuitry are lo- 3.2.2 Discharge Initiation Circuit cated in the boxes behind the main capacitor (see figure 9.) In order to begin the discharge, a small amount of Teflon must be ablated by a smaller, higher voltage 3.2.1 Power Processing and Main Capacitor spark. Control of the spark plugs is handled by the Circuit Fairchild Discharge Initiation (DI) circuit as shown in figure 3. The circuit is duplicated for each of the The 17 µF main capacitor is made of alternating four spark plugs although there are only two 625 V, layers of aluminum, mylar, paper, and mylar filled 2 µF DI capacitors, one each for right and left noz- with 1016 Aroclor Oil. The capacitor is cylindrical in zles. Logic circuitry inside the APPT permits only shape with a diameter of 12.7 cm, a height of 8.9 cm, one spark plug to fire at a time. Each spark plug has and a mass of 1.93 kg. The capacitor was designed its own 3000 V, 0.02 µF capacitor on the secondary and constructed with care taken to provide the lowest side of the transformer. internal inductance possible (<15 nh) while insuring at least a lifetime of 100 billion pulses[3]. The charging circuit uses a flyback type converter which keeps charging power constant. For the COM- 3.3 Chassis Heater Control and PASS P3OINT flight module, the current and voltage Thermal Management supplied to the APPT power processing unit (PPU) was recorded during operation showing that the PPU In both the operational and stand-by modes, the uses 53 W regardless of the voltage. Charging lasts heating system of the APPT module may be switched 0.48 s giving the PPU a total of 25.4 J. During test- on. The heating system has internal control circuitry ing, the 17 µF main capacitor was shown to store and external temperature monitoring equipment de- 17.7 J before the discharge. The pulse efficiency of signed to keep the thruster between 20◦ and 50◦ C. the PPU is, therefore, 70% with the energy stored in The APPT is designed to be thermally− isolated from the spark plug capacitors being neglected. the rest of the spacecraft so that only the APPT The entire circuit is also designed to have the low- heater control system is required for proper ther- est inductance possible. The main capacitor is con- mal management. When power is supplied to the nected directly to both sets of electrodes by two strip thruster and the APPT module temperature drops lines yielding an average inductance value during the below 50◦ C, the heater turns on. Above this tem- discharge of 40 nh. Although the main capacitor is perature, the heater is off. The temperature of the always connected to the electrodes, the circuit cannot APPT module is monitored by a thermistor (WSI be completed and the discharge cannot begin until a Precision Thermistor 44031) mounted on the right small amount of Teflon is ablated by the spark plugs, side capacitor casing as shown in the thruster draw- as described in the next section. ings (see figure 9). ZIEMER, CUBBIN, CHOUEIRI, ORAEVSKY & DOKUKIN: COMPASS P3OINT MISSION 7

4 Power, Command, and Telemetry 2.5 35 Voltage 30

2.0 Voltage (Volts) During the mission, power will be supplied by the 25 batteries and solar panels on board the COMPASS 1.5 20 satellite. Command will be supplied by ground con- 15 trol with computers on the satellite converting the 1.0 10

information to APPT compatible command signals. Current (Amps) 0.5 Five telemetry signals will show the main capacitor 5 Current and spark plug capacitor charging cycle voltages, the 0.0 0 progression of both Teﬂon propellant bars, and the 0.0 0.5 1.0 1.5 2.0 temperature of the main capacitor. More detail for Time (s) each of these topics is presented in this section.

4.1 Power Requirements Figure 4: Voltage and current drawn from the power supply during charging of the APPT capacitors with The APPT requires an unregulated (not current lim- a +27 V supply and the heaters off. ited) power supply providing 15-31 V. Peak current requirements can be above 3 A initially although normal operating conditions may only require 2 A de- rates of up to two pulses per second can be tolerated. pending on the power supply voltage. The APPT has The pulse rate for most of the operational modes of internal circuitry to protect the satellite power supply the flight module will be 0.89 Hz. The capacitors are from an internal short circuit. The results discussed charged in 0.48 s regardless of the power supply volt- in this paper use an unregulated +27 V (+/– 4 V) age or pulse frequency. The average power required power supply as a baseline for the COMPASS Satel- for capacitor charging alone is equal to, lite power bus. 53W 0.48s f . (2) This sub-section of the report includes power re- × × pulse quirements during capacitor charging and stand-by 3 modes as well as heater power requirements. The av- For the COMPASS P OINT Mission, an average erage power required by the APPT module depends power consumption of 25 W is desirable. Allowing on the pulse rate and temperature conditions. With for heater power and control circuit demands, this is the heating system on and operating at a pulse rate of obtained by setting the pulse frequency at 0.89 Hz. one discharge every 1.12 seconds (0.89 Hz) the APPT will use 25 W of power. 4.1.2 Chassis Heater The chassis heater control circuitry draws 5 mA of 4.1.1 Power Processing Unit current whenever power is supplied to the APPT The main discharge capacitor and the spark plug ini- module. When the heater is activated, an additional tiation capacitors are charged simultaneously at the 83 mA (at 27 V) is required to power a 325 Ω resistor. beginning of the pulse sequence. During charging, the The heating system requires a total of 2.38 W when power supply remains at a constant voltage (+27 V) the heater is on and only 0.14 W when the heater is with a current of approximately 1.9 A for slightly less off with a power supply voltage of 27 V. than 500 ms. At a power supply voltage of +23 V the APPT draws approximately 2.3 A. At a power sup- 4.2 Command Signals ply voltage of +31 V the APPT draws approximately 1.7 A. Regardless of power supply voltage, the APPT The APPT requires two digital command signals, one uses approximately 53 W of power during charging. for each nozzle, to begin the pulsing sequence on ei- The 27 V signal is shown, as measured, graphically in ther nozzle. The command signal is a standard 5 V figure 4. The average power consumed by the APPT TTL signal dropping to ground (0 V) to begin the depends on the pulse rate. For the current and volt- pulsing sequence. Only one of the nozzles will be age measurement experiments, our rate of operation commanded to fire at any time; the other will receive was 1.2 seconds per shot (0.83 Hz) although faster a constant 5 V signal. A measured command signal 8 ZIEMER, CUBBIN, CHOUEIRI, ORAEVSKY & DOKUKIN: COMPASS P3OINT MISSION

gression of the right and left teflon propellant bars. 5 The end of the teflon bars have a metal plate that joins two resistive strips. The resistance in the com- 4 Command pleted circuit increases linearly with the progression Signal of the teflon bar. The resistance varies from 80 Ω 3 when no shots have been fired to 6 kΩ when all the 2 propellant has been used. The resistance of both pro- 1500 V pellant strips will be monitored by an external circuit. Telemetry 1 (1 V = 300 V) Since the teflon bars will move a very small amount Signal Voltage (Volts) per pulse, this resistance will be measured only once 0 every 200 pulses with a resolution of 1 Ω. 0.0 0.5 1.0 1.5 2.0 The final signal will come from monitoring one of Time (s) the two telemetry thermistors located on the exterior of the main capacitor. The resistance of the thermistor varies as a function of temperature. At -20 C Figure 5: Trigger signal history shown along with the ◦ the resistance is 78 kΩ. At 50 C the resistance is 1500 V telemetry output from the APPT module. ◦ 3900 Ω. The resistance of the telemetry thermistor will be monitored by an external circuit. along with a sample telemetry signal from the 1530 V capacitor are shown in figure 5. The pulsing sequence begins with the 0 V com- 5 Operation Modes mand signal lasting approximately 15 ms. At that During the mission, there will be various scenarios point, the main discharge capacitor and one of the for operating the thruster. These operational modes two discharge initiation capacitors begin to charge. In are defined here. A table of operational modes is also approximately 0.5 s, both the main capacitor and the included at the end of this paper in figure 12. All spark plug capacitors reach their required voltages, “right” or “left” side conventions are shown on the 1530 V and 635 V, respectively. During this time, no APPT module drawings. other external command signals will be followed. The voltages on the capacitors are monitored by internal circuitry that, when charging is completed, com- 5.1 Power off mands an SCR to close and fire one of the spark plugs In this mode, no power is supplied to the APPT mod- on the appropriate side of the thruster. There are ule. The control circuitry will be off, the heaters will four spark plugs, one on the top and bottom half of not have power, and no command or telemetry signals each nozzle. When one nozzle is commanded to fire, will be monitored. the top and bottom spark plugs of that nozzle will alternate automatically for even propellant ablation. 5.2 Stand-by 4.3 Telemetry Signals In this mode, the APPT is supplied +27 V from the satellite power bus. The power conditioning circuitry For the COMPASS P3OINT Mission, the APPT re- and the chassis heater control will be activated us- quires five analog telemetry signals. Two of the ing a total of 0.68 W. The heating system will turn telemetry signals will monitor the 1530 V capacitor on when the APPT module temperature drops below and the 635 V capacitor voltages. These signals will 50◦C. The heating system uses an extra 2.24 W of be sampled at a minimum rate of 1000 samples per power when the heater is on. Only the thruster tem- second to insure the maximum voltage on the capac- perature will be monitored on telemetry channels. itor charging cycle is recorded. The range for both telemetry signals is 0-5 V. For the 1530 V capaci- 5.3 Mode Rs: Right single pulse tor, the voltage is reduced by a 300:1 voltage divider. For the 635 V capacitors, the voltage is reduced by In this mode, the APPT will be activated for a sin- a 125:1 voltage divider. Both reductions are made gle pulse of the right nozzle. During the charging internally. cycle lasting approximately 500 ms, the APPT will Two other telemetry channels will monitor the pro- use close to 53 W of power. Heating circuitry will ZIEMER, CUBBIN, CHOUEIRI, ORAEVSKY & DOKUKIN: COMPASS P3OINT MISSION 9

also be activated if temperatures drop below 50◦C. Both capacitor voltage telemetries and the thruster From DI Spark Plug temperature will be recorded during this mode. Circuit is shorted (PLUG) to ground 5.4 Mode Rm: Right multiple pulse 17 µF Main Electrodes In this mode, the APPT will be activated for multi- Capacitor ple pulses of the right nozzle at a set rate, 0.89 Hz, to match satellite power supply requirements (25 W with heaters on.) The power conditioning circuitry 5 kW 100 W and the chassis heater control will be activated using High Voltage Relay a total of 0.68 W during off-charging periods. Both capacitor voltage telemetries, the resistance of the right Teflon bar strip, and the thruster temperature Relay closes with +12 V will be recorded during this mode. Safety Switch 5.5 Mode Ra: Right automatic pulse Relay Control In this mode, the APPT will be activated for mul- OR Circuit tiple pulses of the right nozzle at an automatic rate Gate set internally by the APPT power conditioning cir- Trigger Signals +12 V +5 V cuitry. This rate corresponds to the maximum al- from 4 DI Circuits Supply Supply lowable pulse frequency, 2.08 Hz. Accordingly, the capacitors will be continually charged and discharged using 53 W of power on average (contingent on COM- Figure 6: External hardware for relay discharge cir- PASS power supply duty cycle and limitations.) The cuit. chassis heater control will also be activated using 0.14 W at 27 V. Although it is not expected in this mode, if the heating system turns on, it will use an All telemetry channels will be recorded in this mode extra 2.24 W of power. Both 1530 V and 635 V ca- of operation. pacitor telemetry signals, the resistance of the right Teflon bar strip, and the thermistor temperature will be recorded in this mode of operation. 6 APPT Module Preparation and Integration Test 5.6 Mode Ls, Lm, and La The integration of the LES 8/9 APPT with the Com- These three modes: left single pulse, left multiple pass Point Satellite includes a final test at atmo- pulse, and left automatic pulse are exactly similar to spheric pressure of all flight-ready electrical compo- the right single pulse, right multiple pulse, and right nents. The LES 8/9 APPT was designed to operate 5 automatic pulse modes, respectively, described in the at pressures below 10− Torr and cannot fully func- previous sections. All pulse commands will be sent to tion at atmospheric pressure. Therefore, the flight the left nozzle only for all of these operational modes. module circuitry had to be temporarily modified for the test. The requirements for the modifications 5.7 Mode Am: Alternate multiple were: to make the test pulse replicate the real dis- pulse charge as much as possible, to make the test safe for the operators, and to insure the module could return In this mode, the APPT will be activated for multiple to normal operating condition after the modifications pulses alternating between the right and left nozzles. were removed. All three requirements have been met Command signals will be sent to both nozzles at a fre- by installing an external high voltage relay that closes quency of .45 Hz. Average power consumption with when the command signal is sent to the discharge ini- heaters on will still have a maximum value of 25 W. tiation circuitry. A diagram of the external hardware The two signals must be 180◦ out of phase to insure is shown in figure 6. The relay is connected in se- that only one nozzle is commanded to fire at a time. ries with a 5 kΩ resistance (maximum allowed power 10 ZIEMER, CUBBIN, CHOUEIRI, ORAEVSKY & DOKUKIN: COMPASS P3OINT MISSION

100 W) and the 17 µF main capacitor of the APPT where the mass bit, mbit is defined as the amount of module. The entire package including the relay, re- mass ablated per shot andu ¯e is the effective exhaust sistor, and control circuit is mounted in a small box velocity. The LES 8/9 APPT uses a capacitor for on the right side of the thruster. The spark plugs its energy storage device so the stored energy is a of the discharge initiation circuitry are shorted inter- function of the capacitance, C, and the initial voltage, nally, and a cover has been placed over the nozzles to V0, insure safe operation. Finally, the modifications have 1 2 Estored = CV . (5) been kept as simple as possible for easy removal after 2 0 the integration test. The efficiency, η, for the APPT is then, The relay control circuit is connected to the four E I2 spark plug TTL control lines, a safety switch, a 12 V η beam = bit . (6) E m CV 2 and 5 V power supply, and the relay coil. The control ≡ stored bit 0 lines are taken off of the main logic board input to The efficiency has been found experimentally by the discharge initiation (DI) circuitry. When any of measuring these quantities. Full details can be found the four spark plug command lines increase to 5 V in another paper at this conference[10]. A brief ex- triggering the spark plug to fire, the relay is closed planation will also be given here. The impulse bit discharging the main capacitor through the 5 kΩ re- is measured by mounting the APPT on a horizontal sistor. A safety switch can also be turned on at any swing-arm thrust stand and monitoring the velocity time to close the relay and discharge the capacitor. of the arm before and after the discharge. The veloc- The control circuit is powered by a 5 V supply with ity is found through the derivative of a position mea- the relay requiring 12 V to close. surement given by a laser interferometric proximeter For this test, the spark plugs are shorted to ground system (IPS). Individual impulse bit measurements for safety reasons. This will not allow them to spark can be made[9] although for these tests, an average and no other discharges will take place at atmospheric value was used. In fact, all quantities are taken over pressure. During the test, the main capacitor will an average of 300-400 pulses to allow a significant reach its normal value of 1500 V and, therefore, re- amount of mass to be ablated for accurate measure- quires the electrodes to be covered by a plexiglass ment. The mass bit is determined by weighing the shield. All wires to the capacitors and DI circuitry teflon propellant bar before and after the series of are internal to the APPT module. The relay control pulses. Repeated removal and replacing of the bar circuitry is compact and mounted inside a shielded showed no significant erosion. An accurate count of metal box that has been attached to the module. The the number of pulses is kept to determine the mass wires carrying the spark plug trigger signals to the bit. The capacitance is assumed to be 17 µF and control circuit are each twisted with a ground wire the initial voltage is given from the 1530 V telemetry to insure shielding from EMI and eliminate false sig- signal. nals. All additions, both internal and external, can A series of 436 thruster pulses produced the follow- be removed with minimal effort. ing results: an average impulse bit of 285 µN-s 1.7% and an average mass bit of 34.77 µg 0.7%. With± an initial voltage of 1420 V 1% and a± capacitance of 7 Efficiency of the APPT 17 µF, the thrust efficiency± of the APPT module is 6.8%. The efficiency of any electric propulsion device is important for scaling relations and mission plan- ning as well as providing a measure for improvement 8 Conclusions with new ideas and innovations. For pulsed plasma thrusters, the thrust efficiency is defined as the ratio Preparation has been made for the October, 1996 of the average kinetic energy of the exhaust beam to space flight of an LES 8/9 APPT module. The the energy stored before the discharge. The kinetic thruster will be launched on board IZMIRAN’s energy in the exhaust beam is usually defined by us- COMPASS satellite with the experiment being called 3 ing the impulse bit, Ibit, as follows: Mission COMPASS P OINT. The flight module has been described here including general characteris- I = m u¯ , (3) bit bit e tics and information regarding power, command, and 2 1 2 Ibit telemetry. Aspects of thermal control have been ad- Ebeam = mbitu¯e = , (4) 2 2mbit dressed and operational modes during the mission ZIEMER, CUBBIN, CHOUEIRI, ORAEVSKY & DOKUKIN: COMPASS P3OINT MISSION 11 have been defined. To prepare for pre-flight integra- [9] E.A. Cubbin, J.K. Ziemer, E.Y. Choueiri, and tion, the moments of inertia and the center of gravity R.G. Jahn. Laser interferometry for pulsed of the APPT module were found, a circuit permitting plasma thruster performance measurement. In an atmospheric electrical test was developed, and the 24rd International Electric Propulsion Confer- performance of the flight module was measured. The ence, Moscow, Russia, September 1995. IEPC module is now ready for the system integration meet- 95-195. ing in August. [10] E.A. Cubbin, J.K. Ziemer, E.Y. Choueiri, and R.G. Jahn. Pulsed thrust measurements using Acknowledgments The authors wish to thank the laser interferometry. In 32nd Joint Propulsion following people for their valuable time and contri- Conference, Lake Buena Vista, Florida, July bution: George Miller of EPPDyL, Robert Vondra of 1996. AIAA 96-2985. the Air Force, and Richard Sullivan of MIT Lincoln Laboratory. [11] T.R. Kane and Gan-Tai Tseng. Dynamics of the bifilar pendulum. International Journal of References Mechanical Science, 1967. Vol. 9, p.83-96. [12] G.H. Jones. A bifilar moment of inertia facility. [1] R.M. Myers. Electromagnetic propulsion for In 23rd National Conference of the Society of spacecraft. In Aerospace Design Conference, Aeronautical Weight Engineering, Dallas, May Irvine, California, February 16-19 1993. AIAA 18-21 1964. N66-37974, CASI 66N37984. 93-1086. [13] P.B. de Selding. Missle will launch aboard a sub. [2] E.Y. Choueiri. Optimization of ablative pulsed Space News, 7(20), May 20-26 1996. plasma thrusters for stationkeeping missions. Journal of Spacecraft and Rockets, 33(1):96–100, [14] W.E. Wiesel. Spaceflight dynamics. McGraw- 1996. Hill, 1989. [3] R.J. Vondra and K.I. Thomassen. Flight quali- fied pulsed electric thruster for satellite control. Journal of Spacecraft and Rockets, 1974. 11. [4] R.J. Vondra, K.I. Thomassen, and A. Solbes. Analysis of solid teflon pulsed plasma thruster. Journal of Spacecraft and Rockets, 1970. 7. [5] A. Solbes and R.J. Vondra. Performance study of a solid fuel pulsed electric microthruster. In 9th Electric Propulsion Conference, Bethesda, Maryland, April 17-19 1972. AIAA 72-458. [6] W.J. Guman. Solid propellant pulsed plasma micro-thruster studies. In 6th Aerospace Sci- ences Meeting, New York, New York, January 22-24 1968. AIAA 68-85. [7] W.J. Guman, R.J. Vondra, and K.I. Thomassen. Pulsed plasma propulsion system studies. In 8th Electric Propulsion Conference, Stanford, Cali- fornia, August 31-September 2 1970. AIAA 70- 1148. [8] W.L. Ebert, S.J. Kowal, and R.F. Sloan. Op- erational Nova spacecraft teflon pulsed plasma thruster system. In 25th Joint Propulsion Con- ference, Monterey, California, July 10-12 1989. AIAA 89-2497. 12 ZIEMER, CUBBIN, CHOUEIRI, ORAEVSKY & DOKUKIN: COMPASS P3OINT MISSION

162.0

193.0

244.1

Figure 7: Drawing number 1, front view of APPT. ZIEMER, CUBBIN, CHOUEIRI, ORAEVSKY & DOKUKIN: COMPASS P3OINT MISSION 13

a 33.2

43.5

15.3

26.9

CL 120 o

ALL DIMENSIONS ARE IN MILLIMETERS

Figure 8: Drawing number 2, bottom view of APPT. 14 ZIEMER, CUBBIN, CHOUEIRI, ORAEVSKY & DOKUKIN: COMPASS P3OINT MISSION

31.0

6.0

123.5 107.5

Figure 9: Drawing number 3, side view of APPT. ZIEMER, CUBBIN, CHOUEIRI, ORAEVSKY & DOKUKIN: COMPASS P3OINT MISSION 15

152.40

114.30 5 holes @ 76.20 D=4.00 38.10 CL

Z-AXIS

NOTE:XYZ AXIS ORIGIN IS AT THE CENTER OF GRAVITY

NOTE: CENTER OF GRAVITY LOCATED ON CENTER LINE CG Y-AXIS

KEY

TEFLON FEED TUBES 152.40

ALL DIMENSIONS ARE 167.5 IN MILLIMETERS

205.3

Figure 10: Drawing number 4, back view of APPT. 16 ZIEMER, CUBBIN, CHOUEIRI, ORAEVSKY & DOKUKIN: COMPASS P3OINT MISSION

115

19.7

Figure 11: Drawing number 5, top view of APPT. ZIEMER, CUBBIN, CHOUEIRI, ORAEVSKY & DOKUKIN: COMPASS P3OINT MISSION 17

Figure 12: Table of APPT Operational Modes 18 ZIEMER, CUBBIN, CHOUEIRI, ORAEVSKY & DOKUKIN: COMPASS P3OINT MISSION

expression for the moment of inertia about the vertical axis in terms of the period and other measurable x values. L A.2 Simplified Model Figure 13 shows the simplified bifilar pendulum configuration. This model assumes that the axis of ro- R tation passes vertically through the center of mass and that it is equidistant from the two wires. In this case, the wires have the same tension equal to one half of the weight (W ) of the body. The wires provide a restoring force whenever the rigid body is rotated about the x axis. The sum of moments (Mx) from the wires, for small displacements, is:

ΣM = W Rsin(θ) W Rθ. (7) x ≈

Figure 13: A simpliﬁed biﬁlar pendulum The desired moment of inertia about the x axis (Ixx) can be introduced using Newton’s Law, ¨ Appendix W Rθ = Ixxφ. (8) Using the geometric relation

A The Bifilar Method for Lθ = Rφ (9) Measuring the Moment of an equation of motion for the θ dimension can be Inertia written: 2 ¨ WR The center of mass and moment of inertia properties θ = θ. (10) LIxx of the APPT were needed for the design and integra- This simple harmonic oscillator system has a natural tion of the COMPASS Satellite. A method known frequency (ω) given by as the bifilar pendulum technique was used to determine these properties. A bifilar pendulum consists WR2 of a rigid body suspended by two wires or filaments ω = . (11) as shown in figure 13. The dynamical equations of s LIxx motion in the general case are nonlinear and com- plicated. The equations of motion can be greatly Rearranging eq.. 11, using D = 2R, and writing τ simplified provided that certain assumptions are not for the natural period gives an expression for Ixx in violated. In this case, knowing the natural twisting terms all measurable quantities: period of the pendulum system about the vertical axis WD2τ 2 allows for the calculation of the moment of inertia of Ixx = . (12) 16π2L the body about this axis. Performing this experiment about six sufficiently different axis of the rigid body A.3 Torsional Correction allows one to obtain the full moment of inertia tensor. A correction to Eq.. 12 must be made to account for the torsional effects of the wires. Modifying Eq.8 to account for this additional torque gives A.1 Equations of Motion GJ The non-linear bifilar pendulum equations of motion W Rθ + 2 φ = I φ¨ (13) L xx are most thoroughly treated in ref [11]. However, if care is taken, a simplified model will accurately de- where G is the shear modulus and J is the polar mo- scribe the bifilar pendulum dynamics and yields an ment of inertia of the wire. The expression GJ/L ZIEMER, CUBBIN, CHOUEIRI, ORAEVSKY & DOKUKIN: COMPASS P3OINT MISSION 19

is the twisting spring constant keff of the wire according to strain theory. Combining Eq.. 9 and 13 zo xe and solving the harmonic system gives the correction term for Ixx as

GJτ 2 k τ 2 o I = = eff (14) x torsion 2π2L 2π2 which agrees with ref.[12]. The final equation used o for Ixx is y WD2τ 2 k τ2 I = + eff (15) xx 16π2L 2π2 Figure 14: The LES 8/9 APPT reference frame and A.3.1 Torsional Spring Constant Measure- a sample x axis orientation. ment keff can be measured experimentally by suspending B.2 Selecting the Necessary Inde- an object from one of the wires and measuring the torsional period. By knowing the moment of inertia pendent Measurements (Itest) of the test body, the harmonic relation Six orientations of the thruster need to be selected. Each experiment will provide a value for the moment 2π keff e = (16) of inertia Ixx about the x axis as labeled in figure 13 τn r Itest and will provide one equation of the form of eq. 17, and the natural period (τn) by experiment, keff can e o o o Ixx Ixx Ixy Ixz be obtained and used in eq. 14. − − eo 1 o o o eo = (R )− I I I R    yx yy yz  − − − Io Io Io − − − zx zy zz B Moment of Inertia Matrix     (18) where Io represents the moment of inertia tensor of Section A.1 provides a means of determining the mo- the thruster in the desired frame. Figure 14 shows e ment of inertia of the thruster about a vertical axis. the o frame and an x axis for which the moment of In order to construct the full moment of inertia ten- inertia (Ixx) is measured in one experimental config- sor (I), the experiment must be repeated with the uration. For example, a right handed rotation of the o thruster suspended in six different orientations. Since o frame about the y axis (see figure 14) of amount φ o e I is a symmetric matrix, three of the nine elements would align the x axis and the x axis of the experi- are repeated. Once six moments of inertia have been ment. Such a rotation is represented by the rotation measured about six sufficiently different axis on the matrix cosφ 0 sinφ thruster, I can be determined for any coordinate sys- eo tem desired. I is constructed from the measured val- R = 0 1 0 . (19)  sinφ 0 cosφ  ues through the use of a rotation transformation. − Placing this into eq. 18 and writing the expression for e B.1 Rotation Transformation the element Ixx gives 2 o o 2 o e The moment of inertia tensor transforms under a ro- c Ixx 2scIxz + s Izz = Ixx (20) tation according to the following relation[14] − where c and s represent the cosine and sine respec- a ab 1 b ab I = (R )− I R (17) tively of the angle φ. One can see that after three e different values of Ixx have been obtained from exper- ab o o o where R transforms from the a frame to the b frame. iment Ixx,Ixz and Izz can be solved for and another This equation allows a measured moment of inertia measurement in the xz plane would be redundant. In o o about an arbitrary axis to be related to the moments this case, Ixy and Iyy can then be obtained from two of inertia about the desired axes. Several such mea- different measurements of the moment of inertia in o surements must be made to determine the entire mo- the xy plane and lastly Iyz from a measurement in ment of inertia tensor in the desired frame. the yz plane. 20 ZIEMER, CUBBIN, CHOUEIRI, ORAEVSKY & DOKUKIN: COMPASS P3OINT MISSION

All moments of inertia for the ﬂight model have been found using this method. The results are presented in section 3.1.