NASA Technical Memorandum 102065 

AIAA-89-2492 

. 

The NASA Low Thrust 

Propulsion Program 

James R. Stone and Gary L. Bennett 

NASA headquarters 

Ofice of Aeronautics and Space Technology 

Washington, D. C. 

Prepared for the 

25th Joint Propulsion Conference 

cosponsored by the AIAA, ASME, SAE, and ASEE 

Monterey , California, July 10- 12, 1989 

(NASA-Tn-102065) TfIE NSA LOY THRUST N89-26895 PROPULSION PROGRAH {HASA) 31 p CSCL 10B 

Unclas 

G3/20. _,02172663 
THE NASA LOW THRUST PROPULSION PROGRAM 

James R. Stone and Gary L. Bennett 

NASA Headquarters 

Office of Aeronautics and Space Technology 

Washington, DC 20546 

SUMMARY 

c\1 

N 

co 

I 

w 

The NASA OAST Propulsion, Power, and Energy Division supports a low thrust 

propulsion program aimed at providing high performance options to a broad range 

of near-term and far-term missions and vehicles. 

major impact on the mission performance Of essentially all spacecraft and vehicles. 

by low thrust propulsion performance and integration features for Earth-toorbit (ET01 vehicles, Earth-orbit and planetary spacecraft, and large platforms 

in Earth orbit. Major emphases of the program are on low thrust chemical, both 

storables and hydrogen/oxygen; low-power (auxiliary) electric arcjets and 

resistojets; and high-power (primary) electric propulsion, including ion, magnetoplasmadynamic (MPD), and electrodeless concepts. This paper will present 

the major recent accomplishments of the program and discuss their impacts. 

Low thrust propulsion has a 

On-orbi t 1 ifetimes, payloads, and trip times are significantly impacted 

INTRODUCTION 

The identification and development of new propulsion systems is fundamental to the continued effectiveness of NASA. Accordingly, NASA continues ta 

maintain a technology program which spans focused development on Earth-to-orbi t 

(ETO) and chemical transfer propulsion, basic propulsion research and technology (R&T>, and advanced propulsion concepts (ref. 1). Essentially all spacecraft and vehicles incorporate low thrust propulsion systems which contribute a 

significant fraction of the vehicle/spacecraft mass, and the cost of delivering 

mass to low Earth orbit (LEO) is a major fraction of the cost of any mission. 

For example, on the Galileo mission, propellant accounts for 43 percent of the 

total mass and for the more challenging (larger velocity increment) comet rendezvous asteroid flyby (CRAF) mission, the propellant fraction is 76 percent 

(ref. 2). 

gram, which is the focus of this paper, is to provide the basis for increased 

performance and operational life of small thruster systems by developing fundamental understanding in the areas of high-temperature materials, fluid 

physics, electrical and magnetic fields, and ion physics. 

The emphasis of this program is on high-risk, high-payoff technologies 

such as chemical thrusters, arcjets, resistojets, ion engines, and magnetoplasmadynamic (MPD) thrusters. The work is principally being carried out 

through two NASA centers: 

Research Center (NASA Lewis). 

program to advance the technology for low-thrust propulsion systems for a 

broad range of future space platforms, spacecraft, launch, and orbit transfer 

vehicles. 

(2) simplifying mu1 tiengine system architecture; and (3) developing the technology necessary to build and test an MPD thruster capable of steady-state 

operation at power levels and several megawatts. 

The objective of the low-thrust primary and auxilary propulsion prothe Jet Propulsion Laboratory (JPL) and Lewis 

Both centers are working on a complementary 

JPL has been focusing principally on (1) improving ion engine life; 

NASA Lewis has been focusing 
principally on (1) improving chemical systems; (2) research on high-performance 

arcjets; and (3) basic physics and scalability of inert--gas ion thrusters and 

high-power MPD thrusters. There is a close interchange between JPL and NASA 

Lewis, and the two programs both complement and supplement each other. 

i nc 1 uded 

Low 

cles inc 

This paper focuses primarily on progress beyond that reported at the 24th 

Joint Propulsion Conference (ref. 31, but some references to earlier work are 

for clarity and continuity. 

CH EM1 CAL 

thrust chemical 

uding ET0 vehic 

propulsion finds application to a broad class of vehies, Earth-orbital and planetary spacecraft and large 

platforms, including Space Station Freedom, in Earth orbit. 

storable propellants are currently used for on-orbit propulsion, but future 

systems may take advantage of using residual primary propellants (hydrogen or 

hydrocarbons and oxygen) for these functions. 

the use of scavenged residuals and from the increase in specific impulse 

(Isp). 

extended life as well a5 increased performance. A high-temperature thruster 

effort is being conducted using advanced materials and diagnostic techniques 

and is aimed at increased performance and life with state-of-the-art and 

advanced storable propellants. Both storable (monopropellant hydrazine) and 

gaseous hydrogenloxygen (HI01 systems have been investigated for Freedom. 

On ET0 vehicles 

Advantages may accrue from both 

New architectures including propellant resupply create emphasis on 

Due in part to problems with Galileo's 10-N thrusters, interest has been 

neightened in understanding the fundamentals of small rockets. Support for 

small rocket fundamental research is provided by the Advanced Transportation 

Propulsion and Low Thrust Propulsion programs. This effort, which was initiated under the Low Thrust Propulsion program, is discussed briefly herein. 

Storable Propellants 

Low-thrust, storable, monopropellant and bipropellant propulsion is widely 

used on a variety of spacecraft, ranging from small satellites to the space 

shutt 1 e. 

Bipropellant. - Storable bipropellants are used for many Earth-orbital and 

planetary spacecraft. 

past year has been the achievement of an extraordinary increase (-900 K) in 

chamber operating temperature for a 22-N advanced storable rocket using nitrogen tetroKide (N2O4 or NTO) and monomethyl hydrazine (MMH) (ref. 4). The 

thruster, shown operating in figure 1, consists of iridium-coated rhenium fabricated by a chemical vapor deposition (CVD) process wherein the iridium coating is first deposited on an expendable mandrel of molybdenum and then the 

thicker rhenium layer deposited on top of it (ref. 5). The rhenium offers the 

desired high-temperature strength while iridium provides the needed oxidation 

resistance. Because of the high melting point of iridium temperatures 

approaching 2700 K are possible. This can also translate into a greater than 

tenfold increase in the operating life. This technology clearly overcomes the 

problems that existing thrusters have presented to NASA and other spacecraft 

users. 

A major accomplishment at NASA Lewis and Aerojet this 

2 
Aerojet first tested two iridium-coated rhenium thrust chambers with 8:l 

expansion ratio, and over 15 hr of operation was demonstrated at temperatures 

from 2250 to 2530 K without failure or any evidence of corrosive attack or 

measurable change in throat dimensions. Similar results were recently demonstrated on a 22-N thruster with 150:l expansion ratio fabricated from the same 

materials by the same technique (ref. 6). The significance of this can be seen 

in figure 2 which compares the state-of-the art (SOA) disilicide-coated niobium 

alloy chamber with the Ir/Re chamber. 

extended lifetimes but it offers higher Isp. 

advance, NASA initially selected this technology for the Mariner Mark 11 comet 

rendezvous asteroid flyby (CRAF) mission. The reduced propellant mass 

(-600 kg) enabled the addition of three experiments to the spacecraft. 

assure the readiness of this technology for the mission, the NASA Office of 

Space Science and Applications supported Aerojet, under the management of the 

JPL, for fabrication and test of a 440-N iridium-coated rhenium thruster, with 

the CVD processing performed by Ultramet. A 4-hr test indicated feasibility 

of the desired performance and life (ref. 7). While some spalling of the iridium has occurred because of contaminants that affected the iridium deposition 

both engines are serviceable and await additional testing. NASA Lewis is continuing to work with Ultramet to improve the fabrication process. A delay in 

the CRAF mission has recently occurred, resulting in less challenging propulsion requirements and a decision to use a propulsion system provided by NASA's 

European partner in this mission. 

Not only does the Ir/Re chamber offer 

Because of this performance 

To 

Monopropellant. - Although multipropellant resistojets and hydrogen/ 

oxygen were base1 ined for Freedom propulsion, the technology of monopropel lant 

hydrazine propulsion was advanced to provide a design option which might have 

lower development cost (although operating costs would be higher due to the 

requirement for propellant resupply). A life test of a 22-N thruster was conducted under a cooperation program between the Rocket Research Co. and NASA 

Lewis. 

greatest reported for small hydrazine rockets (ref. 8). 

The total impulse capability of 4.8~106 N-s demonstrated is the 

Hydrogen/Oxygen 

Space Station Freedom. - Based on earlier studies manned space stations, 

two propulsion systems were investigated for different functions (refs. 9 

to 11). As discussed elsewhere in this paper, low power multipropellant 

resistojets were selected for drag make-up. for the high-thrust (110 to 220 N) 

function, gaseous H/O thrusters were investigated (refs. 9 to 15). Initial 

thruster designs had an ISp goal of greater than 400 secl and, thus, had 

mixture ratios from 3:l to 5:l. 

for various station fluid management strategies led to the requirement to 

operate at an 8:l mixture ratio with the products of water electrolysis and 

thus reduced the Isp goalto 346 sec. 

design modifications and an experimental demonstration of the 8.9~106 N-s 

design goal. These successful results have led to H/O propulsion being baselined for the high thrust function on the Freedom. 

Subsequent studies of propellant availability 

This led to the need for thruster 

]In this paper Isp is defined as thrust per unit weightflow and has 

units of seconds (sec) in both U.S. customary and '5.1. units. 

3 
Future launch vehicles. - Studies at NASA Lewis have shown that advanced 

auxiliary propulsion subsystems (APS) can significantly increase the payload 

of advanced ET0 vehicles (refs. 16 and 17). The concept is based on using an 

integrated H/O propulsion system in which propellants scavenged from the launch 

vehicle would be used for a liquid hydrogenlliquid oxygen (LH2/L02) primary 

reaction control system (RCS) and a gaseous hydrogenlgaseous oxygen (GH2/G02) 

vernier RCS. Unlike previous studies conducted during the space transportation 

system (STS) technology program, these studies have avoided any propellant conditioning requirements and have suggested the development of hardware to use 

the propellants in the as-delivered state. Figure 3 shows the proposed integrated H/O system in comparison with a 1960's vintage RCS concept. 

The advantages of the currently proposed system include: (1) greater (by 

about 100 sec) system specific impulse; (2) simplified system through elimination of heat exchangerjgasifier, gas accumulator, and compressor; and (3) a 

decoupled thruster/propel lant management system. Figure 4 shows the payoff in 

terms of increased payload mass for the integrated H/O system. A potential 

payload benefit of 4800 kg is projected for Space Shuttle 11. Such a payload 

benefit may also offer significant operational benefits in terms of the types 

of orbits that can be used. An industry study to verify these potential benefits is being initiated by NASA Lewis. 

Rocket Fundamentals 

Present computer codes do not correctly predict the performance of small 

rockets. As a result, NASA Lewis has significantly modified a National Aerospace Plane (NASP) combustor code and extended it to space station H/O rockets. 

This code offers the advantage of including the effects of a shear layer to 

cool the nozzle (for a small rocket the boundary layer is not so thin in relative terms) and of not having numerical "patches" at the throat. The initial 

outputs of the code include combustion shear layer mixing, transonic flow, and 

viscous nozzle flows. 

with NASA's Ames Research Center (ARC) to use the computational capability of 

the National Aerospace Simulator (NAS) for the continuing work (ref. 18). 

The results are encouraging and NASA Lewis will work 

ELECTROTHERMAL 

Electrothermal propulsion using subkilowatt resistojets is currently performing north-south stationkeeping (NSSK) on geosynchronous (GEO) satellites 

(refs. 19 and 20) and will be used for Freedom and man-tended platforms 

(ref. 2). The NASA electrothermal auxiliary propulsion program includes the 

development of resistojet options for high specific impulse and for long duration compatibility with multiple propellants. Kilowatt-class arcjets are being 

developed to provide increased life for higher power geosynchronous satellites. 

Power electronics technology is required for all these concepts and is, therefore, an integral part of the program, as it is also in the high power, primary 

propulsion area. 

Arc je t s 

1-kW preflight development. - NASA Lewis has been studying arc physics; 

electrode materials and their compatibility with propellants; heat transfer, 

4 
plasma arc, and fluid mechanics models. 

included plume fundamentals, arclflow stability, and thruster/power processor 

integration fundamentals. 

in the early 1990's. 

life test of a 1.3-kW arcjet this past year, thereby demonstrating system 

autonomy (ref. 21). No life-limiting mechanisms were identified as a result of 

the test. The arcjet met the pretest goals of 35 percent thrust efficiency and 

450 sec specific impulse. Figure 5 shows the specific impulse as a function 

of the ratio of the power. 

startups of the arcjet (refs. 22 and 23). 

function of cycle number; note that most of the voltage rise occurred in the 

initial 10-hr burn-in period (ref. 21). 

The NASA Lewis experiments have 

NASA is working toward a flight test of an arcjet 

NASA Lewis successfully completed a lOOO-hr, 500-cycle 

NASA Lewis also demonstrated reliable, consistent 

Figure 6 shows the arc voltage as a 

Before arcjet technology can be applied to spacecraft, however, a number 

of spacecraft issues must be addressed. One of these is the thermal impact 

the arcjet will have on both the spacecraft and on the power connection to the 

power processing unit. NASA Lewis has been analyzing the thermal/mechanical 

interface between the arcjet and the spacecraft. The goal is to keep the back 

temperature low enough (~400 K) and the conducted heat low enough (<lo W) to 

minimize problems in designing the rear electrical connector. An obvious 

approach would be to extend the lengths of the arcjet; however, the structural 

analysis showed that the arcjet cannot be made long enough to get the desired 

temperature because of excessive heat losses. 

high emissivity surfaces. Further analyses showed that a coating such as zirconium diboride with an emissivity of about 0.6 would meet the thermal interfaces (ref. 24). 

This pointed out the need for 

Under NASA Lewis contract, supported by the OAST Flight Projects Office, 

Rocket Research Co. has fabricated a flight-type, kilowatt-class arcjet, which 

is shown in figure 7. The electrode configuration has been verified by performance demonstrations and life testing. The arcjet has been successfully 

integrated with the power processing unit (PPU) and has met the thermal interfaces. The dynamic structural analysis of the arcjet has been completed and 

the electromagnetic effects on sate1 1 ite communications have been assessed. 

While many major issues such as electrode life and reliability have been 

addressed (e.g., ref. 21) little has been done to improve the efficiency of the 

device. Preliminary experimental work has shown that the operating characteristics of arcjets depend on the location of the arc attachment zone in the 

anode/nozzle. An anodelnozzle consisting of five pieces of tungsten insulated 

from each other by boron nitride spacers was used in a radiation-cooled, segmented anode to study the physics of the arc attachment zone (ref. 25). Large 

potential gradients were shown to exist in the nozzle during operation with all 

the segments attached to a common point. 

identified by forcing the point of attachment to different positions. Further, 

use of the segments as floating probes gave rough estimates of plasma potential, and, by subtraction, the anode drop. A knowledge of the fundamental of 

arcjets is important because of the existing low efficiency (33 to 38 percent) 

of arcjets caused, in part, because it expends so much energy in breaking bonds 

leading to a "frozen flow" with about 80 to 85 percent of the energy in the gas 

itself. Computationally. nozzle studies are difficult because of the extremely 

small dimensions, low Reynolds number operation, and complicated arc physics 

in the constrictor/nozzle region. 

coupled, which complicates the analyses. 

etry affects the arcjet characteristics (ref. 26). While the performance is 

Preferred modes of operation were 

The plasma and flow fields are strongly 

Figure 8 shows that the nozzle geom5 
insensitive to the upstream design of' the thruster, it appears that a conicaldiverging nozzle is the optimum design. 

As part of its effort to understand the fundamentals of arcjet (and ion 

engine) performance, NASA Lewis has established a laser-induced fluorescence 

(LIF) facility and is now identifying the pliime species as part of its investigation of the electrode losses and spacecraft interactions. Under internal 

funding, JPL will be measuring the velocity and temperature profiles in the 

exhaust stream of a 30-kW arcjet using LIF. 

imental setup (ref. 27). 

Figure 9 shows the JPL LIF experPlume impacts. - As the arcjet thruster moves closer to being a flightready propulsion system onboard communications satellites for applications such 

as NSSK, spacecraft integration concerns become of significant interest. A key 

issue centers on the potential impact that the arcjet plume may have on the 

transmission of RF signals in common use for satellite communications. A first 

order study using Langmuir probe data indicated that the impacts appear negligible in the antenna far field, except for propagation paths near the arcjet 

source (ref. 28). More rigorous numerical techniques are necessary to fully 

define the interaction processes and effects which would occur if the plasma 

were to expand within the antenna near field. Recent results (ref. 29) extend 

the previous work and present a body of systematic experimental data which 

correlate measured plasma properties in the arcjet plume with a wide range of 

thruster operating conditions and constrictor geometries and confirm the 

earlier finding that communications impacts should be negligible. 

5-kW System technoloqy. - Recent interest in arcjets has led to the 

development and life tests of both 1 and 30 kW (e.g., ref. 30) arcjet systems 

for application to GEO satellite stationkeeping and prime propulsion. Orbit 

transfer missions from LEO to GEO might significantly benefit from arcjet systems in the 5-kW power range coupled with the use of high thrust chemical 

hydrazine propulsion to rapidly transverse the Van Allen radiation belts. The 

low thrust, high Isp hydrazine arcjet system would then be used to complete 

the orbit raising and inclination changes required to achieve geosynchronous 

orbit. 

voltage buses since high bus voltage results in low spacecraft power system 

mass. 

Such high-power arcjet systems will likely be energized from highA 5-kW class arcjet system was recently developed at NASA Lewis (ref. 31) 

using a full-bridge, pulse-width-modulated power converter with electrical 

isolation. Previously, a 1-kW class pulse-width modulated parallel converter 

was used. 

parallel converter. The bridge converter power transistor switches operate at 

bus voltage whereas the parallel converter transistor switches must support 

twice the bus voltage. This advantage is partially offset by having two 

switches in series, but lower voltage transistors also have lower ON resistance. Another advantage of the bridge converter is that the power transformer 

is simpler and lower mass than the parallel converter power transformer. The 

laboratory 1-kW, radiation-cooled arcjet used for this work is an extension of 

a design described elsewhere (ref. 21). Specific differences from the 1-kW 

class arcjet include modifications in the cathode, constrictor, and thruster 

heat radiating area. 

For high bus voltages, a bridge converter has advantages over the 

6 
Resistojets 

High performance. - To advance storable propellant resistojets beyond the 

current state-of-the-art 300 sec mission-averaged Isp (ref. 32), several 

efforts have been undertaken. NASA Lewis is working through a four-phase contracted industrial development program to design, fabricate, and characterize 

a hydrazine resistojet thruster capable of about 320 sec Isp. The first 

phase, which was conceptual design and analysis, has been completed and Rocket 

Research was chosen for the design finalization and performance characterization. Currently the program is assessing the tradeoff on immersed heaters 

versus radiatively-coupled heaters. 

longer-life higher-temperature heaters (refs. 33 and 341, arid improved nozzle 

performance (refs. 35 and 361 are being sought. 

test faci 1 i ties for evaluating performance has been demonstrated (ref. 37). 

Improved heat exchanger performance, 

The importance of high quality 

NASA Lewis has been modeling the supersonic viscous flow in a resistojet 

by solving the Navier-Stokes equation in conservative form with a finite-volume 

formulation. A typical solution is shown in figure 10, which gives a plot of 

Mach number contours for an axisymmetric converging-diverging nozzle with a 

20" divergent half-angle, area ratio of 20, and 1086 Reynolds number (ref. 38). 

Multipropellant. - -. As previously mentioned, multipropellant resistojets 

have been baselined for freedom. 

most other propulsion applications in that long life and integration features 

are much more important than performance (refs. 9 to 11). Utilizing freedom 

wastes as propellant minimizes resupply requirement and may eliminate the need 

to return some wastes to Earth. Engineering model thrusters have been built 

and delivered to NASA for testing (ref. 39). Performance characteristics on a 

wide range of possible propellants have been obtained (ref. 40). and long-term 

life tests have been conducted for over 10 000 hr with over 140 deep thermal 

cycles (ref. 41). 

resistojets to operate on high-frequency power (ref. 42). To facilitate the 

integration of this technology on Freedom, propellant and feed system options 

were assessed (ref. 43), and interfaces were defined (ref. 44). The exhaust 

plumes of the thrusters are of concern because of potential effects on sensors 

and experiments and potential attenuation of signals propagating through the 

plume. Analytical and experimental techniques have been developed (ref. 45>, 

and a preliminary assessment made of the effect of nozzle geometry on plume 

characteristics (ref. 46). The exhaust of the engineering model resistojet was 

investigated using rotary pitot probes and a rotating quartz crystal microbalance. The effect of a plume shield has been evaluated and found to be 

small, but the absolute mass flux into the backflow region was very low 

(-7x10-7 g/cm2-sec> (ref. 47). 

The requirements are quite different from 

A power controller was developed to enable the Freedom 

Water. - Interest in water resistojets developed as a spin-off to the 

Freedom multipropellant resistojet development. for man-tended, shuttleserviced platforms such as the proposed industrial space facility (ISF), ease 

and safety of propellant resupply and ground handling operations are critical 

issues, which leads to interest in water as a propellant. 

capable of steam operation already under development for Freedom. water resistojets were baselined for the ISF (ref. 48). 

of a zero-gravity steam generator when it was determined that available steam 

generators, such as that developed for the Manned Orbiting Research Laboratory, did not perform stably and reliably and that operation of the freedom 

With a resistojet 

This necessitated the development 

7 
resistojet with liquid water feed could be quite unstable (ref. 40). 

steam generator was developed after evaluation of a number of boiler concepts 

studied for space power application. Further development of this concept led 

to an integrated vaporizer/resistojet utilizing a single heater (ref. 49). 

This concept is shown in figure 11; the liquid water is swirled to the outer 

wall and boiled by heat radiated from the central heater which also heats the 

resistojet by conduction. 

operates at temperatures below saturation. 

stability and phase separation problems typically encountered in zero-gravity 

boilers. 

of conditions, and flow visualization results obtained (ref. 50). Initial 

studies of the plume characteristics of steam resistojets has recently been 

conducted (ref. 51). Mass flux dnd flow angle measurements were obtained in 

the plume using a variable temperature quartz crystal microbalance. 

A cyclone 

Thermal losses are reduced since the outer water 

This design also overcomes the many 

Laboratory models have been built and performance mapped over a range 

PRIMARY (HIGH POWER) ELECTRIC 

Prospects now appear good fcjr broad acceptance and application of electric 

propulsion systems (ref. 20). However, current applications are limited to 

low-power stationkeeping and orbit control (refs. 19, and 52 to 54). The successful operation of these systems has led to increased user confidence, and 

further operational use of electric propulsion appears imminent. 

The National Commission on Space advocated a number of challenging missions, such as a return to the Moon, unmanned and manned exploration of Mars 

and its moons, and unmanned scientific exploration of the rest of the solar 

system (ref. 55). Many of these missions would be enhanced, and some would be 

enabled by high ISp electric propulsion. To perform the challenging future 

missions, high power and high ISp systems will be required. Candidate systems include electrostatic (ion) and MPD engines, with electrodeless approaches 

representing a longer term possibility. Electric propulsion is also applicable 

to advanced robotic exploration missions, offering both decreased propellant 

mass (and/or increased payload) and reduced trip times for deep space missions. 

Furthermore, this technology may be applicable to manned spacecraft using artificial gravity, where speed is not so critical. There are also many potential 

applications for high Isp electric propulsion at the tens of kilowatts available from large solar power systems. High specific impulse clearly offers 

propellant mass savings, but in order to practically exploit that benefit for 

space missions, it is necessary for the overall vehicle to exhibit acceleration 

levels sufficient to meet mission time lines. This requirement necessitates 

low specific mass and high efficiency propulsion and power systems in order to 

keep power system mass low. 

Mission and System Analyses 

Mission analyses covering spacecraft from small, so ar electric propulsion (SEP) (refs. 2 and 56) to large, megawatt nuclear e ectric propulsion 

(NEP) spacecraft (refs. 57 to 61) show enhancing to enab ing levels of capability for exploration of the planets and beyond (ref. 52). JPL considered a 

small, 180-kg demonstration spacecraft with a 1.5-kWe solar array using ion 

propulsion for transportation from LEO to lunar orbit (ref. 56). This Lunar 

Get Away Special (LGAS) mission would carry one instrument, nominally an Apollo 

gamma ray spectrometer, and would carry 36 kg of xenon propellant. JPL has 

8 
also studied a deep space mission concept, wherein the Thousand Astronomical 

Unit (TAU) Explorer, propelled by NEP, would reach 1000 A.U. in 50 years with 

a significant science payload (ref. 57). Both of these spacecraft accomplish 

the mission propulsion with a minimum of propellant mass and, thereby, maximize the science payloads at their destinations. 

(ref. 60) of the benefits ion propulsion might provide to the CRAF mission if 

it were used instead of storable bipropellants. Even though the mission and 

its science package were not optimized for use of electric propulsion, a 3-year 

trip time reduction could be achieved along with a 600-kg increase in launch 

mass margin. 

A study was conducted by JPL 

A range of solar- and nuclear-powered Earth orbital missions have been 

studied and show costjmass reductions benefits (refs. 62 to 64). In a USAF 

study it was shown that a solar electric orbit transfer vehicle for delivery 

of satellite constellations to their operational orbits could offer significant 

savings (ref. 62). 

Nuclear electric propulsion system technology for an electric cargo vehicle has been identified as a major requirement for evolutionary solar system 

exploration (ref. 65). The use of an electric, low-thrust cargo vehicle in the 

Lunar Outpost to Early Mars Evolution case could potentially reduce the initial 

mass in LEO (IMLEO) by one-third over a standard chemically propelled vehicle. 

Studies have shown that for a Mars cargo vehicle large enough to support a 

manned mission, high performance electric propulsion with specific impulses 

over 4000 sec at multimegawatt power levels can offer major reductions in propellant requirements at acceptable transit times. Compared with an aerobraking 

cargo vehicle using hydrogen/oxygen propulsion, and nonaerobraking high performance electric vehicle can result in elimination of at least three Heavy 

Lift Launch Vehicle (HLLV) launches, and savings are even greater when compared 

with nonaerobraking chemical propulsion (ref. 66). This reduction in the number of launches might permit the mission to be completed more rapidly if HLLV 

launch rates are low. To perform these challenging future missions, the total 

impulse capability must be advanced from the currently demonstrated 106 N-s 

for both ion and MPD engines to at least 108 N-s (ref. 67). 

Ion 

Ion engines have demonstrated specific impulses from less than 2000 sec 

Proto more than 10 000 sec, thrust efficiencies to over 0.75 (e.g., ref. 681, and 

total thrust impulse as high as 106 N-s for 10-kW class thrusters. 

pellants have included mercury, argon, krypton, and xenon. Ion thruster technology has been demonstrated with flight testing of the Space Electric Rocket 

Test (SERT 1, 11) ion thrusters (ref. 69). Electrostatic (ion) propulsion is 

planned for its first operational use on the Japanese Engineering Test Satellite VI (ref. 54). 

(fig. 12). But to advance to significantly higher power, the key issues are 

scale-up of ion acceleration subsystems for high power operation and thruster 

life. High power operation is llmited by ion optics, and extraction area must 

be increased with increasing power. Ion propulsion systems in the range of 10 

to 50-kW offer significant payload and trip time advantages over chemical systems for lunar ferry and outer planet explorer missions (including multiple 

target missions), as well as for missions outside the solar system A nearterm target for planetary and Earth orbital missions is about 5 to 30 kW per 

ion thruster using xenon as a propellant. 

Thruster size is adjustable to fit mission requirements 

The planned Pathfinder Cargo Vehicle 

9 
Propiilsion program will focus on a target of 0.1 to 1 MW per ion thruster using 

argon or krypton. While the performance engineering of ion engines is well in 

hand the scalability, system, lifetime, and power processing issues remain open 

and are being addressed in the NASA program. 

Lifetime. - Interest in near-term application of ion thrusters for solar 

electric propulsion in near-Earth missions (refs. 62 and 70) has led to a 

program to identify and extend the operating limits of ion thrusters originally 

developed for 3-kW operation with mercury to the IO-kW range using inert gas 

propellants (refs. 71 and 72). The original mercury ion engines presented life 

limiting phenomena like spalling of metal flakes eroded from the discharge 

chamber surfaces which could short out the screen accelerating grids as well as 

the problem of using a toxic, condensible propellant. As a result the program 

is now focused on inert gas propellants, primarily xenon. JPL modified the 

0.3-m J-series mercury ion engines, which had previously been highly developed 

and were ready for space qualification testing at 2.2 kW, to use xenon. Recognizing that the xenon sputter rates are greater than for mercury and the xenon 

engine operating powers are 2 to 5 times those for mercury engines, lifetime is 

a critical issue. An ion thruster system for planetary applications has been 

designed and is undergoing engineering tests (ref. 73). The focus of these 

activities is on reducing system complexity, with the goal of reducing the 

cost of development and flight qualification. Proposed design changes are 

evaluated in terms of mission impact (ref. 74). 

To overcome cathode erosion problems, JPL has been te;ting a 1.27-cm diameter hollow cathode shown in figure 13 (ref. 75). The cathode was successfully 

operated for 1200 hr on argon, of which 1050 hr were at 100 A and 24 hr were at 

150 A. The cathode was still operating when the testing was terminated; cracking along the tantalum barrel at the upstream end and tungsten deposits which 

were in the process of closing off the insert bore were two of the problems 

observed in operating these cathodes at high power. 

cathode jet formed which could have been responsible for the high erosion rate 

at the upstream side of the baffle. 

showed that the introduction of nitrogen into the xenon propellant reduced the 

baffle erosion. 

grids. The work will be continued (ref. 76). In addition, alternatives to 

tantalum will be tested as baffle materials and erosion and deposition rates on 

critical surfaces and deposit structure will be characterized. 

There may have been a 

Preliminary data obtained by JPL in 1987 

Nitrogen may also reduce the erosion of the accelerating 

Scaling issues. - Results have previously been reported for 30- to 50-cm 

diameter xenon thrusters at power levels up to 20 kW. However, the only 

extended operation of a high.power xenon ion thruster has been with a 30-cm 

J-series divergent-field thruster at 10 kW. This thruster had insufficient 

life for high total impulse mission because of cathode baffle erosion. 

Lewis has addressed its cathode-baffle erosion, which was running about 

0.5 mm/hr, by eliminating the baffleipole piece and going to the ring cusp 

design shown in figure 14. This design will require the application of a magnetic field on the order of 500 to 1000 G at the boundary (refs. 77 and 78). 

Measurements made by NASA Lewis have shown that the erosion rates are very 

sensitive to the discharge current. In order to understand erosion better, 

NASA Lewis is developing a model of sputtering. 

NASA 

NASA Lewis has demonstrated the technology to fabricate 0.5-m ion thrusters (ref. 79) as shown in figure 15. Figure 16 shows the ion beam extraction 

capability of a 0.3-m ion engine and the new 0.5-m ion engine. The 0.5-m 

10 
engine doubles the power capability and extends the specific impulse capability. 

circuit. ) 

(The thruster efficiency is nearly independent of size and magnetic 

Early ion thruster grids (ion optics) were flat and were subject to 

severe, uncontrolled buckling when radial temperature gradients became too 

steep. The grids were then dished to control buckling. Increasing the thrust 

of ion engines will require increasing grid size. With the larger grids and 

potentially noncircular geometries now under consideration, computer modeling 

has been developed and experimentally verified for understanding structural 

response (ref. 80). 

Engine system technoloqy. - JPL has been working on a simplified system 

desiqn arid improving throttling strategy for planetary mission. As shown in 

f i gu;e 17, JPL has determi ned that ion-engi ne; operati ng at f i xed propel 1 ant 

flow rates can be throttled over an input power range of 3.2:l with only a 

small increase in propellant consumed and small decrease in payload delivered 

as compared to operating the engines at a variable flow rate. The improved 

throttling strategy seeks to select the most optimum number of operating 

engines, operational duty cycles, and specific impulse. The systems studies 

have shown that this optimized throttling strategy results in a reduction in 

the required engine lifetime of 25 percent to perform the base mission, which 

was a rendezvous with Comet Kopff, when compared to the conventional throttling technique. Obviously the constant flow strategy eliminates the flow 

controllers and permits operation at fixed discharge conditions - and it may 

also simplify the PPU. JPL plans to configure the laboratory PPU for a twoengine operation and operate the two-engine module on a single screen supply. 

Computer control algorithms for multiengine control will be developed. 

NASA Lewis has examined in parametric fashion the impact of various 

thruster technologies and operating conditions and system technologies on the 

design and performance of a 10-kW class ion propulsion system for orbital 

maneuvering and transfer. A methodology is described wherein a computer code 

is used for precise determination of ion thruster performance (based upon predefined and user-defined thruster technologies), a determinat on of overall 

propulsion system design including a complete mass breakdown. and first-order 

calculations of total impulse and delta-V capabilities. Resu ts are obtained 

from this code to examine the impact of thruster technologies (engine diameter, 

2- and 3-grid ion optics, thruster operating conditions (power level propellant, and specific impulse)), and subsystem technologies on system design and 

mission capability. 

(ref. 81). 

These analyses were also conducted for arcjet systems 

MAGNETOPLASMADYNAMIC 

Multimegawatt MPD thruster technology offers payload gains of a factor of 

10 over chemical propulsion for typical orbit raising, lunar base ferry, and 

manned interplanetary missions. The MPD thruster combines high specific 

impulse (1000 to 5000 sec) with moderate thrust levels (10 to 100 N). Historically, steady state MPD thrusters have typically been run at or below the few 

hundred kilowatt range (e.g., refs. 82 to 84), while MPD thrusters operated in 

a pulsed mode have been run at power levels exceeding 1 MW (e.g., refs. 84 

to 87). Propellants have included H2, Ar, He, Li, Ne, "3, and N2H4. 

11 
a 

wi 

de 

High specific impulse (3000 to 6000 sec) at efficienc es of 30 to 50 percent 

have been demonstrated using hydrogen (e.g., refs. 88 and 891, though in general thrust efficiencies have not exceeded 30 percent for most propellants 

(fig. 18). Presently, MPD thruster lifetimes are lim ted by severe cathode 

erosion (refs. 90). MPD propulsion is less advanced han ion. Higher efficiencies and specific impulses are generally obtained with applied magnetic 

fields, but the fundamental theoretical understanding of this mode of operation 

is lacking. 

25 kW (ref. 91). 

The highest total impulse demonstrated is 106 N-s, at about 

Currently, NASA Lewis is involved in both experimental and theoretical 

research in the areas of cathode erosion and MPD thruster performance. NASA 

Lewis has made operational its high-power MPD thruster test stand and successfully demonstrated an operating power of 130 kW in a hybrid MPD/electrothermal 

mode and GO kW in the MPD mode (ref. 91) NASA Lewi(; has extended the magnetic 

field to 0.3 T and shown reduced anode damage on start-up and an extended 

stable steady-state operating range as shown in figure 19 (refs. 92 and 93). 

A high-power MPD thrust stand (ref. 94) and nonintrusive plasma diagnostics 

(ref. 95) are being developed, and hardware is being assembled for a series of 

cathode erosion experiments. Codework is being written to simulate the magnetoplasmadynamic interactions (ref. 96), and will be used to evaluate thruster 

design changes (ref. 90). JPL is also working on such a test stand and both 

Centers are developing nonintrusive diagnostics for MPD testing. The potential 

impact of high-temperature superconductivity developments on MPD technology 

was assessed in a joint effort by Argonne National Laboratory and NASA Lewis 

(ref. 97). 

Work at JPL has demonstrated that cathode evaporation can be minimized by 

designing for thermionic emission during steady-state operation. This may lead 

to practical thruster lifetimes (21000 hr). In the near term, JPL is focusing 

on demonstrating MPD thruster operation for periods greater than 100 hr. Previously, using the thruster concept shown in figure 20, JPL had completed 85 

runs totaling 13 hr of operation, reaching a maximum power of over 72 kW at 

2245 A. The engine operated for an 83-min period with no excessive arc spot 

damage on the cathode. The cathode tip temperature was less than 2240 K for 

operation up to 1700 A with an argon propellant mass flow rate of 0.16 g/sec. 

JPL has observed two distinct steady-state operating modes as shown in 

figure 21. The first mode is characterized by a luminous cathode jet while the 

second mode has a cooler cathode tip and no cathode jet. At the same current 

the terminal voltage is as much as 40 percent lower in the second mode. It may 

turn out that the second mode could provide the more efricient acceleration of 

the plasma, due to a more contained radial current pattern. Work is continuing 

to understand the conditions under which the two modes occur. 

As part of its effort to improve MPD thruster performance JPL has designed 

radiation-cooled, applied field hybrid engine as shown in figure 22. This 

11 be fabricated and tested. JDL has also been reviewing the literature to 

termine if designs previously discarded because of the previously existing 

ectrode sheaths 

ons and percated twoMPD thrustertechnology limitations may now- be useful. Recognizing that e 

have a significant effect on MPD thruster operational limitat 

formance, (JPL is sponsoring the development of a more sophist 

dimensional model that will provide a better understanding of 

operations than current one-dimensional models. 

12 
ELECTRODELESS DISCHARGE THRUSTERS 

Electrode thruster concepts offer the potential for very high energy coupling efficiency, and are therefore, of great interest for high power applications (refs. 98 to 104). Limitations imposed by electrode erosion in other 

electric propulsion thruster designs are potentially avoided. 

Currently NASA Lewis is evaluating the microwave electrothermal (MET) 

thruster (ref. go), which absorbs applied microwave power in a plasma discharge 

and heats the gas propellant by high pressure thermalization (fig. 23). Advantages include a high efficiency for power absorption; a high power density; 

and external control of the discharge location, shape, and volume. The MET 

thruster should readily scale with power (a1 though performance may not) and it 

is compatible with high power operation. The current technology MET, being 

developed in conjunction with Michigan State University, operates at a microwave frequency of 2450 MHz and a power coupling efficiency in excess of 95 percent. Power levels up to 2.5 kW have been tested, and specific impulse values 

up to 600 sec have been obtained with helium. Recent results show energy efficiencies approaching 65 percent at low power (ref. 105). Gases tested include 

He, N2, H2, and 02, and a variety of microwave applicator designs have 

been demonstrated. Present limitations in performance are due to material 

constraints and MET throat/nozzle design (ref. 90). 

Future work at NASA Lewis will focus on power levels of 30 kW at a frequency of 915 MHz, using H2, He, and N2 at discharge pressure up to 10 atm. A 

magnetic nozzle, produced by a low temperature superconducting magnetic magnet 

(peak field 5.7 TI, will be used to increase the plasma density in the discharge chamber walls. Computer modeling will be used to predict performance 

with magnetic nozzles, and to compare performance 1 imi ts with other electrodeless concepts including electron cyclotron resonance (ECR) and ion cyclotron 

resonance (ICR) heated thrusters (refs. 90, 99, 105, and 106)). JPL is also 

conducting research on the ECR concept (ref. 104). 

ELECTROMAGNETIC LAUNCHERS 

The possible application of electromagnetic launchers to deliver nonfragile materials from the Moon to LEO or other nodes was suggested by the 

National Commission on Space (ref. 55). Several years ago NASA Lewis completed an investigation of electromagnetic launchers for delivery of nonfragile 

payloads from Earth into various orbits (refs. 107 to 109). These studies 

indicated significant cost benefits for Earth-based launchers if the launch 

rate of appropriate cargo is high enough. Studies have also considered the 

utility of EML's for the launch of raw materials and nonfragile cargo from the 

lunar surface (refs. 110 to 113). However the launch rates would have to be 

higher than those envisioned in any of the exploration case studies before an 

Earth-to-space rail launcher would be competitive with chemical launchers. 

Advances in high temperature superconductivity could contribute significantly 

to the utility of such launchers (ref. 114) and make them the most economical 

means of bulk transport from the surface of the Moon for such materials as 

propellants (ref. 115). 

13 
CONCLUDING REMARKS 

The ongoing NASA program in low-thrust primary and auxiliary propulsion 

continues to make great strides. The development of high-temperature chemical 

thrusters offers order-of-magnitude lifetime improvements and higher specific 

i!iipulses. 

the are baselined for space station. Work is advancing on arcjets leading to 

a p anned flight demonstration. 

are being addressed with the goal to flight demonstrate these systems so they 

wil be ready for future missions such as a cargo vehicle flight to Mars. 

The technology for resistojets has been taken to the point where 

Both ion and MPD thruster technology issues 

ACKNOWLEDGMENTS 

The authors acknowledge the contributions made by P. Garrison, J. Kelley, 

D. King, T. Pivirotto, J. Barnett, K. Goodfellow, W. Deininger, and C. Garner 

of JPL and by D. Byers, F. Curran, R. Myers, S. Schneider, J. Sovey, and 

L. Carney of NASA Lewis. 

of G. Reck and E. VanLandingham of NASA Headquarters which allowed us to complete this paper. 

The authors also acknowledge the continued support 

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Devices," Proposed for the 1989 JANNAF Propulsion Meeting, Cleveland, OH, 

May 1989. 

Examination," AIAA Paper 89-2832, July 1989. 

95. Myers, R.M., "Plume Characteristics of MPD Thrusters: A Preliminary 

96. LaPointe, M.R., "Numerical Simulation of Applied Field Steady State MPD 

Thrusters," proposed for the 1989 JANNAF Propulsion Meeting, Cleveland, 

OH, May 1989. 

97. Reed, C.B., and Sovey, J.S., "The Use of High Temperature Superconductors 

in Magnetoplasmadynamic Systems," NASA TM-101219, 1988. 

98. Pollard, J.E., Lichtin, D.A., and Cohen, R.B., "RF Discharge Electrothermal Propulsion: Results for a Lab-Scale Thruster," AIAA Paper 

87-2124, June 1987. 

99. Whitehair, S., Asmussen, J., and Nakanshi, S., "Microwave Electrothermal 

Thruster Performance on Helium Gas," Journal of Propulsion and Power, 

Vol. 3, No. 2, Mar.-Apr. 1987, pp. 136-144. 

100. Morin, T.J., and Hawley, M.C., "The Efficacy of Heating Low-Pressure H2 in 

a Microwave Discharge," Plasma Chemistry and Plasma Processinq, Vol. 7, 

No. 2, June 1987, pp. 181-199. 

20 
101. Hawley, M.C., Asmussen, J., Filpus, J.W. Frasch, L.L., Whitehair, S., 

Morin, T.J., and Chapman, R., "A Review of Research and Development on the 

Microwave-Plasma Electrothermal Rocket," AIAA Paper 87-1011, May 1987. 

102. Frasch, L.L., Griffin, J.M., and Asmussen, J., "An Analysis of Electromagnetic Coupling and Eigenfrequencies for Microwave Electrothermal 

Thruster Discharges," AIAA Paper 87-1012, May 1987. 

103. Filpus, J.W., and Hawley, M.C., "A Computer Model for the Recombination 

Zone of a Microwave-Plasma Electrothermal Rocket," AIAA Paper 87-1014, 

May 1987. 

104. Sercel, J., "Electron-Cyclotron-Resonance (ECR) Plasma Thruster Research," 

AIAA Paper 88-2916, July 1988. 

105. Whitehair, S., and Asmussen, J., "The Experimental Performance of a Compact Microwave Electrothermal Thruster," Michigan State University, East 

Lansing, MI, 1988 (to be published). 

106. Whitehair, S., Frasch, L.L., and Asmussen, J., "Experimental Performance 

of a Microwave Electrothermal Thruster with High Temperature Nozzle Materials," AIAA Paper 87-1016, 1987. 

107. Zana, L.M., Kerslake, W.R., and Sturman, J.L., "Rail Accelerators for 

Space Transportation - An Experimental Investigation," NASA TP-2571, 1986. 

108. Rice, E.E., Miller, L.A., and Earhart, R.W., "Preliminary Feasibility 

Assessment for Earth-to-Space Electromagnetic Rai lgun Launchers ,I' NASA 

CR-167886, 1982. 

109. Miller, L.A., Rice, E.E., Earhart, R.W., and Conlon, R.A., "Preliminary 

Analysis of Space Mission Applications for Electromagnetic Launchers," 

NASA CR-174748, 1 984. 

110. Kolm, H., "An Electromagnetic Slingshot for Space Propulsion," Lechnology 

111. O'Neil, G.K., and O'Leary, B., eds., Space Based Manufacturing from NonReview, Vol. 79, June 1977, pp. 60-66. 

terrestrial Materials, Progress in Astronautics and Aeronautics, Vol. 57, 

AIAA, New York, 1977. 

112. Heppenheimer, T.A., "Achromatic Trajectories and Lunar Material Transport 

for. Space Colonization." Journal of-Spacecraft and Rockets, Vole 15. No. 

3, May-June 1978, pp. 176-183. 

113. Billingham, J., Gilbreath, W.P., O'Leary, B. , and Gosset, B., eds., Space 

Resources and Space Settlements, NASA SP-428, NASA, Washington, DC, 1979. 

114. Hull, J.R., and Carney, L.M., "Application of Superconducting Technology 

to Earth-to-Orbit Electromagnetic Launch Systems," NASA TM-101134, 1988. 

115. "Exploration Mission/Technology Planning Workshop." Workshop Proceedings, 

Vols. I & 11, Washington, DC, Nov. 14-17, 1988. 

21 
ORIGINAL PAGE 

BLACK AND WHlTE PHOTOGRAPH 

FIGURE 1. - ADVANCED 22-N Bi-PROPELLANT THRUSTER. 

1960’s RCS CONCEPT PROPOSED RCS CONCEPl 

oms 

TANK OMS PRQPhlANl 

TANK CONOlllOWlNG 

GHylGOz RCS THRUSTERS 

PROPELLANT CONDITIONING 

LHz1LOz RCS THRUSTERS 

NO PROPELLANT CONDITIONING 

* ELIMINATED COMPONENTS 

- HEAT EXCHANGERIOASIFIER 

- OAS ACCUMULATOR 

-COMPRESSOR ID ‘1R.31PI4 - OECOUPLEO THRUSTERIPROPELLANT MANAGEMENT 

SOA (Si ON Nb) Ir ON Re 

1000 r (Isp = 280-290s) (Isp = 306-311s) -l 1000 

1500 2000 2500 3000 3500 4000 4500 

CHAMBER TEMPERATURE. OF 

FIGURE 2. - PERFORMANCE AND LIFE EXTENSION 

PROVIDED BY IRIDIUM-COATED RHENIUM THRUST 

CHAMBER, 

SHUTTLE II SORTIE MISSION’ 

50 r 

S 0 A EARlH 

STORABLE 

SYSTEM 

SCAVENGING 

HI0 SYSTEM 

ILn=l27 I~C 

SCAVENOING 

HI0 SYSTEM 

I.”=lll rec 

LOADED 

PROPELUNI 

ORY MASS 

Itp=31068c 

A PAYLOAD BENEFIT OF 10,700 Ib PROJECTED FOR SHUTTLE II 

MAY OFFER SIGNIFICANT OPERATIONAL BENEFITS 

‘VEHICLE MASS AND MISSION REQUIREMENTS SUPPLIED BY LaRC 

cD-BB-11311 

FIGURE 3. - INTEGRATED HI0 AUXILIARY PROPULSION SYSTEM. FIGURE 4. - PAYOFF OF INTEGRATED H/O PROPULSION. 

22 
OptGlMAL PAGE IS 

POUR QUALITY 

47c 

45c U 

v) 

W 

v) 

a -1 

8 43c 

: 

LL 

w 

V W 

410 

250 

0 

MASS FLOW RATE. 

KG/SEC 

o 4.1~10-~ 

0 4.6~10-~ 

0 5.0~10-~ 

OPEN SYMBOLS DENOTE DATA 

CLOSED SYMBOLS DENOTE DATA 

HALF SYMBOLS DENOTE DATA 

THREE QUARTER SYMBOLS 

TAKEN BEFORE LIFE TEST 

TAKEN AFTER 144 CYCLES 

TAKEN AFTER 335 CYCLES 

DENOTE DATA TAKEN AFTER 

LIFE TEST CONCLUSION 

POWER. W 

FIGURE 5. - LIFE-TEST 1-KW ARCJET SPECIFIC 

IMPULSE AS FUNCTION OF POWER. 

250 - 

0- 

FIGURE 6. - LONG-TERM VOLTAGE TRENDS FOR LIFE-TEST 1-KW ARCJET. 

23 
ELECTRODE CONFIGURATION VERIFIED BY: 

- PERFORMANCE DEMONSTRATIONS 

> 

W 

+ :: 100 

9 

- LIFE TESTING 

PPU INTEGRATION DEMONSTRATED 

THERMAL INTERFACES MET i DYNAMIC STRUCTURAL ANALYSIS COMPLETED 

COMMUNICATIONS IMPACTS ASSESSED 

. 

FIGURE 7. - FLIGHT-TYPE, KILOWATT-CLASS ARCJET 

I I I 

150 1 mi BASELINE 

ox ox DOWNSTREAM CONTOUR 

DUAL 

A CONTOUR 

450 r 

400 ' 

U Y v1 

0 

0 

I I 

10 15 20 25 30 

P/M, J/Kg 

FIGURE 8. - NOZZLE GEWETRY EFFECTS ON ARCJET CHARACTER ISTICS. 
PULSED DETECTOR 7 

I TUNABLE NARROWBAND 

INTERFERENCE 

FILTER 

VOLUME COOLED 

VACUUM 

TANK c 1.2X2.1 M 

TO PUMP 

FIGURE 9. - JPL LASER INDUCED FLUORSCENCE (LiF) EXPERIMENTAL 

SETUP. 

A INTERNAL THERMOCOUPLE 

* EXTERNAL THERMOCOUPLE 

LOCATIONS 

LOCAT IONS 

WICH CONTOURS 

AXISYM NOZZLE, AR = 20. 41 x 21 GRID. RE = 1086, DT = .Ole 500 

STEPS 

CONTOURS: 2 = 0.5. 3 = 1.0, 4 = 1.5. 5 = 2.0. 6 = 2.5. 7 = 3.0, 

8 = 3.5. 

FIGURE 10. - COMPUTED R4CH NUMBER CONTOURS FOR LOW REYNOLDS NUMBER CONVERGING-DIVERGING NOZZLE. 

LIQUID 

6 H -FLOW RESTRICTOR 

BOILER CHAMBER /-LIQUID INJECTOR 

SUPERHEATER INSERT 

u 

1 2 cm SUPERHEATER-’ 

FIGURE 11. - INTEGRATED WATER VAPORIZER/RESISTOJET. 

25 
NEAR TERM PLANETARYlEARTH ORBIT MISSION TARGET = 5 TO 30 LWITHRUSTER, Xe PATHFINOER CARGO MISSION TARGET 0.1 TO 1 MWITHRUSTER, At OR Kr 

THRUSTER 

POWER, 

kW 

loo 

lono i 

lJ & 

STATIONKEEPING SEPS 

CIRCA 

1980 

Ar.Kt 

H XE 

SEPS CLASS CARGO 

PLANETARY/ VEHICLE 

EARTH ORBIT PROPULSION 

TRANSFER 

FIGURE 12. - ION ENGINE POWER RANGE FOR 

VARIOUS MISSIONS. 

NEW AFTER 170 HR AT 100 A 

FIGURE 13. - JPL 1.27 CM DIAMETER HOLLOW CATHODE. 

DIVERGENT FIELD RING CUSP 

. Ulolrll 

. 

VOLUME 

BAFFLE FOR DISCHARGE 

IMPEDANCE CONTROL 

USED FOR SERT I, SERT It, IAPS, 

AND SEPS 

BOUNDARY MAGNETIC FIELD 

NO BAFFLE 

FIGURE 14. - DIVERGENT-FIELD AND RING-CUSP ION THRUSTERS. 

26 ORIGINAL PAGE 

BLACK AND WHiTE PI-iOTOCRAPH 
FIGURE 15. - 50 CM ION THRUSTER. 

LARGE AREA ION OPTICS PROVIDE: 

2X INCREASE IN POWER OVER 30 cm DIA. THRUSTER 

EXTENDED SPECIFIC IMPULSE CAPABILITY 

25 - 

XENON 

POWER, 

kW IO5- 

0 I 

2000 3000 4000 5000 

SPECIFIC IMPULSE, s 

FIGURE 16. - ION EXTRACTION CAPABILITIES 

OF 30 cn AND 50 cn THRUSTERS. 

27 

0 CONSTANT FLOW STRATEGY ELIMINATES FLOW 

CONTROLLERS AND PERMITS OPERATION AT FIXED 

DISCHARGE CONDITIONS - MAY SIMPLIFY PPU 

0 ENGINE LIFE REQUIREMENT REDUCED 24% BUT REQUIRES 

9% MORE PROPELLANT 

3000 Kg PAYLOAD 

TITAN IV LAUNCH 

75 KW ARRAY STRATEGY 

HOURS N-S Kg 

REQUIRED TOTAL PROPELLANT 

ENGINE LIFE IMPULSE MASS 

FIGURE 17. - ION ENGINE THROTTLING STRATEGY. 
ORIGINAL PAGE 

BLACK AND W.MjTE Pl-m 

.8 r 0 HYDROGEN 0 FLAG: STEADY STATE 

0 "A" APPLIED FIELD I I!? 0 AMMONIA 

D NITROGEN 

0 HYDRAZINE 

0 ARGON 

u 0 OA U 

A, .4 

I 

0 11111 I I I111111 I I I II lld 

101 102 103 104 

INPUT POWER, KW 

FIGURE 18. - DEMONSTRATED EFFICIENCY FOR MPD THRUSTERS. 

ANODE (+)7 

/ 

I/ LBUFFER ELECTRODE 

, 

FIGURE 20. - JPL STEADY-STATE, SELF-FIELD MPD THRUSTER. 

ARC 

VOLTAGE. V 

ARC PRESSURE - 100 TORR 

MAGNETIC FIELD 

0.1 0.2 

MAGNETIC 02 MAGNETIC FIELD. TESLA 

APPLIED MAONETIC FIELD 

TESLA 01 EXTENOEO TO 0.3 TESLA 

FIELD. "1 / 

REDUCED ANODE DAMAGE ON START-UP 0 5w EXTENDED STABLE STEADV.STATE 

CU-81-37102 ELECTROMAGNET CURRENT, A OPERATING RANGE 

FIGURE 19. - OPERATING RANGE FOR LERC APPLIED-FIELD NPD THRUSTER. 

DISTINCT STEADY-STATE OPERATING MODES 

OBSERVED BY D. KING IN 1987 

ONE MODE IS CHARACTERIZED BY A LUMINOUS 

CATHODE JET. A SECOND MODE HAS A COOLER 

CATHODE TIP AND NO CATHODE JET; THE TERMINAL 

VOLTAGE IS LOWER. 

SECOND MODE MAY PROVIDE MORE EFFICIENT 

PLASMA ACCELERATION 

FIGURE 21. - STEADY-STATE, SELF-FIELD MPD MODE CHANGES. 

28 
-RADIATOR 

ELECTROMAGNET -. 

FIGURE 22. - JPL RADIATION-COOLED, APPLIED-FIELD MPD.THRUSTER 

DESIGN. 

REGENERATIVE y THROAT-NOZZLE ASSEMBLY 

THROAT-NOZZLE 

COOL I NG 7 1F-e 

VIEW I NG 

WINDOW 

CENTER CONDUCTOR-ULv 

-\- PROPELLANT 

I INFLOW (PAR TIAL) 

1 IN. 

FIGURE 23. - ELECTRODELESS. MICROWAVE PLASMA TORCH THRUSTER 

DESIGN. 

29 
Report Documentation Page 

_______-- 

9 Sycurity Classif (of this report) 

__ __ Spice Adin,nislr,i c I 

- _. ___ ~ - - ___ 

1 Report No NASA TM-,02065 7 Government Accession No 

22 Price' 

-_ __ AIAA-89- 2492 _. . 

4 Title and Subtitle 

The N .\SA Low Thrust Piopulsion Program 

_____- 

7 Author@) 

James R. Stone and Gary L. Bennett 

-- -~~~ - 

9 Perforniiny Organization Name and Address 

NASA Headquarters 

Office of Aeronautics and Space Technology 

Washington. D. C. 20546-OOO 1 

2. Sponsoring Agency Name and Address 

National Aeronautics and Space Administration 

Washington. D.C. 20546-0001 

I 

~ 

5. Supplementary Notes 

.- ~- -__-- 3. Recipient's Catalog No 

- _______ 5. Report Date 

6. Performing Organization Code 

_____.-_ 

8 Performing organization Report No 

E-4822 

10. Work Unit No. 

506-42-3 1 

11. Contract or Grant No. 

13. Type of Report and Period Covered 

Technical Memorandum 

14. Sponsoring Agency Cod; 

Prepared for the 25th Joint Propulsion Conference cosponsored by the AIAA, ASME, SAE, and ASEE, 

Monterey, California, July 10-12, 1989. 

The NASA OAST Propulsion, Power, and Energy Division supports a low thrust propulsion program aimed at 

providing high performance options for a broad range of near-term and far-term missions and vehicles. Low 

thrust propulsion has a major impact on the mission performance of essentially all spacecraft and vehicles. Onorbit lifetimes, payloads. and trip times are significantly impacted by low thrust propulsion performance and 

integration features for Earth-to-orbit (ETO) vehicles. Earth-orbit and planetary spacecraft, and large platforms in 

Earth orhit. Major emphases of the program are on low thrust chemical propulsion, both storables and hydrogen/ 

oxygen; low-power (auxiliary) electric arcjets and resistojets; and high-power (primary) electric propulsion, 

including ion, magnetoplasmadynamic (MPD), and electrodeless concepts. This paper will present the major 

recent accomplishments of the program and discuss their impacts. 

7. Key Words (Suggested by Author@)) 

Electric propulsion 

Low thrust 

_____._~ 

Unclas\ified - Unlimited 

Sublcct Category 20 

Unclassified I Unclassified I 30 I A03 

'For sale by the National Technical Information Service, Springfield, Virgiiiia 22161 NASA FORM 1626 OCT 86