r- ~ -- - ~ - - - --- -,-- - -. - - - I NASA Contractor Report 198427 Pulsed Plasma Thruster Technology for Small Satellite Missions Roger M. Myers and Steven R. Oleson NYMA,Inc. Brook Park, Ohio Melissa McGuire Analex Corporation Brook Park, Ohio Nicole J. Meckel and R. Joseph Cassady Olin Aerospace Company Redmond, Washington November 1995 Prepared for Lewis Research Center Under Contract NAS3-27186 National Aeronautics and Space Administration PULSED PLASMA THRUSTER TECHNOLOGY FOR SMALL SATELLITE MISSIONS Roger M. Myers and Steven R. Oleson Nyma Inc. NASA Lewis Research Center Cleveland, OR 44135 Melissa McGuire Analex Corp. NASA Lewis Research Center Cleveland, OH 44135 Nicole J. Meckel and R. Joseph Cassady Olin Aerospace Company Redmond, WA Presented at the 9th AIAAJUtah State University Conference on Small Satellites Sept. 18-21, 1995 . .. ... __ .__ ._-..... PULSED PLASMA THRUSTER TECHNOLOGY FOR SMALL SATELLITE MISSIONS Roger M. Myers and Steven R. Oleson Nyma, Inc. NASA Lewis Research Center Cleveland, OH 44135 Melissa McGuire Analex Corp. NASA Lewis Research Center Cleveland, OH 44135 Nicole J. Meckel and R. Joseph Cassady Olin Aerospace Company Redmond, WA Abstract Pulsed plasma thrusters (PPTs) offer the combined benefits of extremely low average electric power requirements (1 to 150 W), high specific impulse (- 1000 s), and system simplicity derived from the use of an inert solid propellant. Potential applications range from orbit insertion and maintenance of small satellites to attitude control for large geostationary communications satellites. While PPTs have been used operationally on several spacecraft, there has been no new PPT technology development since the early 1970's. As a result of the rapid growth in the small satellite community and the broad range of PPT applications, NASA has initiated a development program with the objective of dramatically reducing the PPT dry mass, increasing PPT performance, and demonstrating a flight ready system by October 1997. This paper presents the results of a series of near-Earth mission studies including both primary and auxiliary propulsion and attitude control functions and reviews the status of NASA's on-going development program. Introduction The continuing emphasis on cost reduction and Pulsed plasma thrusters rely on the Lorentz spacecraft downsizing is forcing increased force generated by the interaction of an arc emphasis on reducing subsystem mass and passing from anode to cathode with the self­ integration costs. For many commercial, induced magnetic fields to accelerate a small scientific, and DoD missions, on-board quantity of ablated chloroflourocarbon propulsion is either the predominant spacecraft propellant. As shown in Figure 1, the thruster mass or limits the spacecraft lifetime. system consists of the accelerating electrodes, Additional pressures resulting from the energy storage unit, power conditioner, ignition emphases on use of smaller launch vehicles, circuit, propellant feed system, and telemetry. new spacecraft architectures, and the costs During operation, the energy storage capacitor is associated with ground testing and handling first charged to between 1.5 and 2.5 kV. The toxic or hypergolic propellants have also led to ignition supply is then activated to generate a the consideration of alternative propulsion low density plasma which permits the energy technologies. The characteristics of pulsed storage capacitor to discharge across the face of plasma thrusters make them uniquely suited for the chloroflourocarbon propellant bar. This arc providing a very simple, light weight, low ablates, heats, and accelerates the propellant to volume, high performance propulsion option for generate thrust. Peak arc current levels are power-limited small satellites. typically between 2 and 15 kA, and the arc 1 --, I I duration is between 5 and 20 ~s. The pulse providing a total impulse of 20,000 N-s. These cycle is repeated at a rate compatible with the objectives will be accomplished via use of available spacecraft power. This ability to use recently developed capacitors, integrated circuit the same thruster over a wide range of technology for both telemetry and power spacecraft power levels without sacrificing electronics, new structural materials, and an performance or having a complex throttling increase in PPT performance. Extensive algorithm is one of the advantages of PPTs. laboratory testing has demonstrated that PPTs The propellant feed system consists solely of a could be built to provide over 2000 s Isp at over negator spring which pushes the solid 20% efficiency) Following completion of the chloroflourocarbon bar against a stop on the initial program, an effort is planned to continue anode electrode, eliminating safety and miniaturizing the PPT if there is sufficient reliability concerns with valves or pressurized interest in the small spacecraft community. systems. There are no other moving parts on the PPT, resulting in a propulsion system which The very low power requirements, small size, is extremely inexpensive to integrate onto and simplicity of integration make PPTs suitable spacecraft and can be stored indefinitely with for a range of small satellite missions. To little concern for storage environment. The demonstrate the potential applications, a set of latter was recently demonstrated when PPTs potential PPT missions were analyzed, stored for over 20 years were successfully fired including orbit maintenance of a small satellite in at both the NASA Lewis Research Center sun-synchronous orbit, drag make-up and orbit (LeRC) and the Olin Aerospace Company raising for a shuttle launched small satellite, (OAC). The largest mass components of the orbit raising and drag make-up for the spacecraft PPT are the energy storage unit (a capacitor or in the Teledesic constellation, and momentum pulse-forming network) and the system wheel replacement for attitude control of both electronics, including the power conditioning large and small satellites. Following a unit, discharge initiation, and logic and description of these mission examples, NASA's telemetry circuits. Recent developments in these PPT development program is summarized technologies provide several options which can including the contracted flight system result in a system mass reduction by a factor of development program, efforts to explore more two. advanced PPT options, and studies of spacecraft integration requirements. PPTs were extensively developed in the late 1960's and early 1970's. Figure 2 shows the range of impulse bits demonstrated on flight or flight-qualified systems. The PPT system Mission Application Examples developed during that period with the most flight experience was used on the Navy's The PPT system assumed for the mission TIP/NOV A navigation satellites and operated at application studies provided either 1000 or a peak power level of 30 watts during firing. 1500 s Isp at an efficiency of 15% and was The NOV A PPT had a specific impulse (Isp) of capable of providing 20,000 N-s total impulse. 543 s, an impulse bit of 400 ~N-s , a total A fueled system mass of 3.5 kg was used, with impulse capability of 2450 N-s, and a fueled operating power levels between 1 and 150 W. system mass of 6.8 kg.l The baseline Throttling is achieved by varying the pulse technology for the new NASA program is the frequency at a constant performance level. flight-qualified LES 8/9 PPT system, which These assumptions are consistent with the was selected because of its higher Isp of 1000 s objectives of NASA's PPT flight system and demonstrated total impulse capability of development program. 10,500 N-s.2 The LES 8/9 operated at power levels of 25 or 50 W, produced an impulse bit of 300 ~N-s, and had a fueled system mass of Orbit Maintenance 6.7 kg.2 As discussed in detail below, the Two orbit maintenance missions were examined. The first, maintenance of a 100 kg initial NASA program objectives are to decrease 2 the fueled system mass to 3.5 kg while satellite with a 0.38 m cross-section in a sun- I I 2 l -- --.---~---- -------..--.' - - - """----=-,- --- - synchronous orbit for five years, had a total the orbit if the mission starts from a higher mission velocity change requirement of 122 initial orbit. This is clearly illustrated by m/s. This mission, if performed by a standard considering a 50 kg satellite with a cross­ hydrazine monopropellant propulsion system, sectional area of 0.43 m2 launched from the would require a propulsion system dry mass of Shuttle at altitudes between 375 and 425 km. 19.3 kg with 7 kg of propellant. The Worst case atmospheric drag was again monopro~ellant system would have a volume of assumed. For these cases, a constant PPT 0.022 m . To perform the same mission, a power level of 30 W was assumed with the PPT system would weigh 8.4 kg (two thrusters) thruster either operating only when the arrays and carry 1.36 kg of propellant, requiring a were illuminated or continuously throughout the volume of 0.012 m3. The PPT would consume orbit. As discussed above, the latter would and average of less than 5 W of electrical power require that the spacecraft power system include during the mission. The 18 kg propulsion batteries capable of providing 30 W throughout system mass reduction achieved using PPTs the shaded period of the orbit. The 30 W power could be used to increase payload mass or, if level limited the PPT thrust level to 0.92 rnN. multiple satellites were launched on a single Results are shown in Table 1, from which it can launch vehicle, the total launch mass savings be seen that the PPTs can raise the spacecraft would result in a substantial increase in initial orbit for either the continuous or illuminated­ delivered altitude. This increase could be over only thrusting cases. The long transfer times 300 km if eight spacecraft were launched on a could be reduced by increasing the power level Pegasus XL. should it be available.