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Spacecraft Propulsion Technology BACKGROUND AND STATUS Propulsion systems on board spacecraft perform orbit transfer, attitude pointing and control, orbit altitude maintenance, north-south or east-west station keeping in geosynchronous orbits, orbit raising from low Earth orbits up to and including geosynchronous Earth orbit, and in-space primary propulsion. Each maneuver places an emphasis on various performance characteristics of the propulsion system, such as thrust level and specific impulse, and not all missions require the propulsion unit to perform all of the cited operations. However, the propulsion system must be capable of operating in various modes to meet the needs of the mission. These modes range from individual engine pulses (possibly for station keeping) to long-duration, steady-state thrusting (perhaps for interplanetary missions). In addition, if clusters of small spacecraft are used for missions requiring simultaneity of measurements, a propulsion system with very high accuracy and precision may be required for station keeping. The smaller mass, moments of inertia, and volume of the small spacecraft drive the desired characteristics of the propulsion system. For on-orbit operations of small spacecraft, the thrust levels must be smaller than those on large spacecraft to keep the acceleration levels within the design limits. The impulse bits cielivere~i for pulsed operation also must be smaller to allow the spacecraft to stay within the bounds of the stabilization control logic. Additionally, the propulsion system volume and weight must be minimized. Spacecraft maneuvers are usually done with chemical propulsion systems. Past improvements of these systems, while impressive, have been incremental. In the future, dramatic improvements in propulsion technology for small spacecraft could be accomplished through other types of propulsion such as electric propulsion. For example, the low weight of the small spacecraft opens up the orbit-change operation to electric propulsion. The very low-thrust electric propulsion devices can be employed for payload placement in orbit or on interplanetary trajectories within reasonable time frames and can possibly reduce the size of the launch vehicle required. 23

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24 Technology for Small Spacecraft CHEMICAL PROPULSION Chemical spacecraft propulsion devices, other than those user! for orbit elevation or orbit insertion, usually employ liquid reactants as the energy source. The propellant might be a single reactant (monopropelIant) or a combination of fuel and oxidizer (bipropelIant) . The most common monopropellant is hydrazine. It is passed through a catalyst bed, where it decomposes into ammonia and nitrogen at a temperature of about 700°C with a delivered specific impulses of about 230 seconds. A monopropellant propulsion system is relatively simple and is amenable to short, pulsed operation, which is suitable for small spacecraft attitude control. The most common bipropeliant system utilizes a nitrogen tetroxide oxidizer and a fuel of either hydrazine or monomethy! hydrazine. The reactants are hypergolic,2 facilitating ignition under vacuum conditions and pulsed operation. Use of hypergolic reactants also provides the capability to restart the system when necessary. The delivered specific impulse of such systems is about 310 seconds. A chemical energy propulsion system integrated into a typical spacecraft operating in Earth orbit can range from 10 to 20 percent of the total spacecraft weight; or up to 40 to 50 percent if significant parts of higher orbit insertion ant! circularization are included in its mission cycle. Technology advancements have therefore focused on achieving higher specific impulse since 90 percent of the propulsion system weight is usually propellant. Most thruster technology focuses on increasing the allowable operating temperature and duty cycle life of radiation-cooled combustion chambers, and on achievement of very small reproducible impulse bits without major degradation in specific impulse. In some special applications with minimal total impulse requirements, thruster weight may be an important factor and some research and development has been focused on thruster weight reduction. For monopropellant systems using hydrazine, research has been focused on increasing the pulse duty-cycle life by reducing catalyst bed degradation, and on increasing the specific impulse by use of electric energy for raising the decomposition products temperature before expansion through the nozzle. NASA Chemical Propulsion Programs The Lewis Research Center (LeRC) is working with Aerojet General, TRW, Atlantic Research, and Ultramet to develop high-temperature, oxidation-resistant materials for small, bipropellant, high specific impulse rockets across a broad spectrum of thrust levels (22 to 550 newtons). The designs incorporate a rhenium-iridium thrust ' Specific impulse is the impulse delivered to the spacecraft per unit weight of expelled propellant. 2 Hypergolic substances are ones that react spontaneously upon contact.

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Spacecraft Propulsion Technology chamber that can permit operation at temperatures up to 2200°C. To date, LeRC has designed, fabricated, and tested four different rockets. One system has tentatively been baselined on an advanced commercial communication spacecraft. Specific goals of the project include attaining specific impulses greater than 350 seconds, a factor of three or greater reduction in rocket sizes anti masses, and the ability to operate radiation-cooled rockets at arbitrary propellant mixture ratios with all on-board propellant options (Bennett, 19941. While not necessarily an enabling technology for small spacecraft in general, these thrusters offer a higher performance, chemical propulsion option for orbit raising, while retaining payload delivery time to orbit on the order of hours. Department of Defense Chemical Propulsion Programs , 25 Much of the recent, low-thrust, propulsion technology that has been developed within BMDO programs can be used as a technology base for small NASA spacecraft. The primary issues become those of capitalizing on the potential for large weight savings, extending the life of the propulsion system for scientific rather than military requirements, and tailoring the size of the system to small spacecraft applications. The issues also include finding techniques for increasing thrust chamber reliable lifetimes to permit the long operating times demanded by some scientific missions. Technology advances and extensive reductions in weight ant! size of pulsed, bipropelIant and monopropellant chemical propulsion units have been made with the recent work on kinetic-kill vehicles by BMDO under the Light Exo-Atmospheric Projectile program. Uncler this program, a 755-newton bipropellant thruster weighing 64 grams, developed by the Rocket~yne Division of Rockwell International, has been flown on a prototype kinetic-kii! vehicle at the Air Force's National Hover Test Facility at EdwarcIs Air Force Base, California. The thruster's rapid-response valving system makes it suitable for numerous spacecraft maneuvering functions. Concurrent with these developments, monopropellant attitude control thrusters (223 to 500 newtons) developed by Rocket Research Company with high pulsing rates and low weight (~84 to 326 grams) have also been demonstrated on the kinetic-kill vehicle. Advances in lightweight, piston-pump, propellant supply systems have been made by Lawrence Livermore National laboratory in conjunction with Rocket Research Company and Moog Valve. These advances reduce feed system weight by about 50 percent while improving engine performance (Whitehead, 19931. The Advanced Liquid Axial Stage program funded by BMDO has also made significant advances in reducing the propulsion system weight while demonstrating the practicality of using carbon composites for spacecraft structures. Carbon composites have been used as the high-temperature thrust chamber structure for high-performance small thrusters and in the construction of high-pressure, bipropelIant propulsion tanks. Use of a carbon-fiber overwrap on a thin aluminum tank liner permits high-pressure tankage operation at one-half the weight of presently used tanks. Under Air Force sponsorship through the U.S. Air Force Phillips Laboratory and the Space Systems Division, Aerojet General Corporation and Rocket~yne have

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26 Technology for Small SpacecraJi demonstrated 16,680-newton orbit transfer engines that use a nitrogen tetroxide and monomethy! hydrazine bipropeliant to deliver a specific impulse of 340 seconds at the time of this study. These engines, designated XER-132s, represent a major upgrade in the technology and enhance payload delivery capability. At the time of this study, the Aerojet General Corporation engine has amassed 680 seconds of operating time at simulated altitude; the Rocket~yne version has accumulated 700 seconds of test time primarily at sea level. SOLAR PROPULSION Whereas chemical propulsion devices use the energy of chemical reactants, solar propulsion devices use the sun's energy to generate high-temperature gases that are expelled at high velocities from a thruster. There are two methods by which this is accomplished. First, a solar electric thruster can convert the solar energy to electrical energy by means of solar cells. This electrical energy is then used to power a thruster in which the electrical energy is converted to the kinetic energy of the expelled, high- temperature gases. The second method captures the solar energy in the cavity of a solar thermal thruster, where its thermal content is absorbed by a working fluid that, in turn, is expelled for thrusting purposes. Solar Electric Propulsion Solar electric propulsion is a near-term technology with considerable potential for reducing spacecraft mass and cost. Electric propulsion generally is characterized by low thrust and high specific impulse. While it cannot satisfy requirements for prompt cleolovment at high altitudes. it is well-suited for less urgent. anticipated demands. r - - r ~- - on _ . . ~ J ~ ~ Electric propulsion can result in spacecraft weight reduction that could dramatically reduce costs by allowing the selection of a smaller launch vehicle. It also can reduce or eliminate the use of gravity assists in planetary missions by enabling direct trajectories, as well as shorter trip times. Even missions to the outer planets court! utilize the solar electric propulsion for continuous thrusting out to perhaps three astronomical units before solar flux diminishes beyond a useful intensity. For surveillance or remote sensing missions that require frequent maneuvering or repositioning, the comparatively high efficiency of electric thrusters can substantially increase spacecraft lifetime or enhance versatility. I ow-thrust, electric propulsion collie also be well-suited for precision station keeping of clusters of small spacecraft. The three basic types of electrically powered thrusters, in or(ler of successively greater potential for higher specific impulse are arc jets, electromagnetic (plasma) thrusters, and ion engines. Arc jets can deliver a specific impulse from about 450 to 550 seconds, whereas electromagnetic thrusters can provide a specific impulse in the range of 1,000 to 2,000 seconds, and the xenon ion engine can deliver a specific impulse in the range of 2,500 to 3,500 seconds. The power required to be delivered to the propulsion

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Spacecrap Propulsion Technology system will be proportional to the operating specific impulse, the thrust profile required to satisfy the mission and the overall efficiency of the unit. Depending on the application and system, design, power levels for small spacecraft could range from ten's of watts to many kilowatts. Power requirements for electric propulsion include both the power delivered to the spacecraft and to the energy residing in the ejected propellants. This total power is proportional to the thrust times the operating specific impulse. The optimization of any size spacecraft for venous missions where electric propulsion may be advantageous requires careful trade offs between power level, specific impulse, ant! thrust profiles. These optimizations will usually determine the selection of arc jets, plasmas, or ion-type thrusters and their operating profiles during the mission. Such profiles can be very different for orbit-raising velocity increments versus station-keeping impulse bits or versus the energy input profile for interplanetary flights where trajectory plans wouic! dominate. However, power requirements for orbit-change velocity increments and interplanetary flights generally will require higher power levels. Recently U.S. private industry has begun to use small arc jet propulsion systems for station keeping of geostationary communications satellites (Aerospace America, 1993; Space Technology Innovation, 1994~. However, there are no flight-qualified single thruster modules currently available at power levels of one kilowatt or more (note: the "desirable" power level is totally mission and device oriented). The xenon ion engine has a higher specific impulse than the electromagnetic and arc jet thrusters. To date, no commercial spacecraft manufacturer has flown ion propulsion engines. However, Hughes has baselined a xenon ion propulsion system for the HS601 Galaxy large spacecraft scheduled for launch in 1995 which court! be applicable to small spacecraft. The xenon thruster at the 1-kilowatt class power level for small spacecraft would require work to improve the ion optics of the thruster, which impacts thruster life. Ton thruster technology is well-suited for interplanetary science missions using Delta Il.- class launch vehicles, small spacecraft (100 to 300 kilograms), and short mission durations. This technology could also be directly applicable to commercial spacecraft for station keeping. NASA Programs for Solar Electric Propulsion Development of the arc jet and ion thruster technologies are underway at both NASA LeRC and the U. S. Air Force Phillips Laboratory. In addition, Boeing is currently exploring a solar electric propulsion option for the Pluto Fast Flyby mission with internal funds. While a variety of technical issues still need to be resolved, studies on the Pluto Fast Flyby mission inclicate that a solar electric propulsion unit that utilizes three high voltage, 5-kilowatt-electric xenon ion thrusters may enable a Delta-ciass or AtIas-cIass launch vehicle to be used instead of a Titan IV. The electric power for these thrusters is generated by solar panels, which employ mini-dome concentrators to enhance electrical output of the solar (photovoTtaic) cells. Such a unit would make trip times of less than ~ ~ years possible through the use of Earth-gravity-assist maneuvers compared 27

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28 Technology for Small Spacecraft to the baselined 8-year direct flight without gravity assists. Use of solar electric propulsion for the Pluto Fast Flyby mission deliberately pushes technology and would require significant investments in design, development, test, and evaluation to reach acceptable levels of mission risk. NASA is scheduled to flight qualify a 2.5-kilowatt xenon ion engine aboard the U.S. Air Force Phillips Laboratory's Electric Insertion Transfer Experiment, which is tentatively scheduled for launch in 1998. DoD Programs for Solar Electric Propulsion BMDO is sponsoring research and development work on the SPT-70 and SPT-IOO electromagnetic Hall thrusters from Russia. Both devices are in the power range of interest for orbit raising and other functions for small spacecraft. Currently, tests are being conducted at LeRC and JPL to validate the performance and life of the thrusters and determine the radiation fields generated about the thruster during operation (Space News, 19931. Preliminary ground testing has produced mixed results clue to the highly ionized exhaust plume, which can interfere with communications, and the wide plume divergence angle, which means the thrusters must be canted to avoid hitting delicate parts of the spacecraft. BMDO flew the SPT-70 electric thruster on the MST} II spacecraft that was launched in May, 1994 (Mattock et al., 19931. The test flight of the SPT-IOO device was planned in conjunction with the test flight of the Topaz reactor, which, at the time of this report, has been inclef~nitely deferred. Ammonia arc jet thrusters are planned for use in the U.S. Air Force Phillips Laboratory's Electric Propulsion Space Experiment, scheduled for launch in 1995 and the Electric Insertion Transfer Experiment, scheduled for launch in 1998 (Avila, 1992; Sneegas et al., 1993~. Solar Thermal Propulsion Solar thermal propulsion offers promise of a higher thrust capability than electric propulsion at Tower specific impulse. Research indicates that these devices for propulsion may be able to deliver a specific impulse of about 850 seconds with higher thrust levels than solar electric thrusters (less than one newton), but they are less developed. The demonstrated specific impulse to date is on the order of 600 seconds. The solar collector (mirror) technology requires! is dependent upon successful demonstration of lightweight, deployable mirrors that can be easily packaged and then deploye(i in space. These mirrors must have concentration ratios on the or(ler of 10,000: ~ to produce the desired thrust levels for small spacecraft. Both the thruster require(1 to absorb the radiated energy that is needed to heat the working fluid and the mirror collector technologies require significant development efforts before solar thermal propulsion can be fully demonstrated.

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Spacecraft Propulsion Technology NASA Programs for Solar Thermal Propulsion NASA does not have any solar thermal propulsion programs. DoD Programs for Solar Thermal Propulsion Technology for solar thermal propulsion is being supported mainly through the DoD Small Business Innovative Research (SBIR) funds and company-funded efforts. Research and development of solar thermal devices has been underway at the U.S. Air Force Phillips Laboratory and in company-funded efforts in 0.22-newton thrust-level devices at Rocketdyne and Hercules (Pande, 19931. The solar collector technologies needed to implement a solar thermal device also are being pursued through the U.S. Air Force Phillips Laboratory. NUCLEAR PROPULSION Both nuclear electric propulsion and nuclear thermal propulsion have been considered by mission planners (e.g., for the Mars mission studies). These technologies are generally incompatible with the assumption in this report of 600 kilograms as the upper limit for small spacecraft. Nuclear propulsion does, however, have the potential of raising a I,000-kilogram spacecraft from low Earth orbit to geosynchronous~Earth orbit with Atlas-class boosters. The technology, therefore, may be addressed in future Aeronautics and Space Engineering Board studies. FINDINGS AND PRIORITIZED RECOMMENDATIONS The Panel on Small Spacecraft Technology believes that advanced propulsion technology can provide dramatic reductions in the cost of placing the payload in orbit. Specifically, electric propulsion can be an enabling technology for small spacecraft in that its use for orbit-raising functions can effect a reduction in launch vehicle size with an attendant reduction in launch costs. It can also be an enabling technology for small spacecraft cluster station keeping. Miniaturization and weight reduction technologies have been demonstrated on chemical propulsion systems. In turn, these technologies lead to reductions in the mass and volume of propulsion systems aboard the spacecraft, which are necessary for on-orbit functions such as attitude control, repositioning, and station keeping. However lifetimes of months to years are required for such applications and necessitate additional technology advancement to ensure suitability of the demonstrated technologies. The panel recommends that NASA focus its technology development and integration resources, first, on solar electric propulsion technology for primary (orbit transfer stage or spacecraft) propulsion; second, on advanced chemical propulsion; and 29

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30 Technology for Small Spacecraft third, on solar thermal propulsion technology. Specific recommendations on propulsion technology are prioritized below. I. An aggressive program should be established to demonstrate, in ground tests, the life of xenon ion propulsion systems that operate at power levels in the range from about 0.5 kilowatt to about 2.5 kilowatts for lifetimes of up to 8,000 hours. Arc jet thrusters for small spacecraft applications also should be evaluated. The systems demonstrated should be capable of being integrated into solar electric propulsion systems with tote] power levels in the range of ~ to 5 kilowatts. Both the ion thruster and the arc jet should then be demonstrated in space flight tests in the near term. A. ~ The propulsion system requirements should be determined for precision station keeping of clusters of small spacecraft, and the capability of currently available systems should be evaluated. If it is necessary, systems should be clevelopec! to meet specific mission requirements. 3. A technology program should be established to demonstrate the [Light Exo- Atmosohenc Proiectile nronulsion technolo~ie^s nt mission ~ ~ ~ ~ duty cycles and lifetimes consistent with small spacecraft mission life and operational requirements. 4. The 445-newton rhenium-iridium thruster should be evaluated for application to an apogee kick stage for small spacecraft. This includes demonstration over a duty cycle typical of the missions envisioned for small spacecraft. 5. The suitability of the XER- ~ 32 engine as an upper-stage propulsion system for launching small spacecraft with creep space propulsion needs should be evaluated. 6. Research and technology programs should be initiated to demonstrate fully the capability of solar thermal rockets, with emphasis on concentrator/mirror, absorber- thruster, and feed-system technology. Space flight tests should be conducted to explore deployment mechanisms and dynamics, validate packaging techniques, and demonstrate the performance and durability of absorber-thruster operation with a deployable concentrator mirror.