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NASA Technical Memorandum 102321

Development of a High Power Microwave

Thruster, With a Magnetic Nozzle,

for Space Applications

John L. Power and Randall A. Chapman

Lewis Research Center

Cleveland, Ohio

Prepared for the

24th Microwave Power Symposium

sponsored by the International Microwave Power Institute

Stamford, Connecticut, August 21-23, 1989

_NASA-TM-102]21) DEVELOPMENT OF A HIGH

POWER MICROWAVE TflRUS_ER, WITH A MAGNETIC

NOZZLE, FOR SPACE APPLICATIONS _NASA.

Lewis Research Center) 26 p CSCL 2|H

G3/20

N89-26904

Unclas

0224225

DEVELOPMENT OF A HIGH POWER MICROWAVE THRUSTER, WITH A

MAGNETIC NOZZLE, FOR SPACE APPLICATIONS

John L. Power and Randall A. Chapman*

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio 44135

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I. SUMMARY

This paper describes the current development of a hlgh-power microwave

electrothermal thruster (MET) concept at the NASA Lewis Research Center. Such

a thruster would be employed in space for appllcations such as orbit raising,

orbit maneuverlng, statlon change, and possibly translunar or transplanetary

propulsion of spacecraft.

The MET concept employs low frequency continuous wave (CW) microwave power

to create and continuously pump energy into a flowing propellant gas at relative high pressure via a plasma discharge. The propellant is heated to very

hlgh bulk temperatures while passing through the plasma discharge region and

then is expanded through a throat-nozzle assembly to produce thrust, as in a

conventlona] rocket engine.

Apparatus, which ]s described, is being assembled at NASA Lewis to test

the MET concept to CW power levels of 30 kW at a frequency of 915 MHz. The

microwave energy is applied in a resonant cavity appllcator and ]s absorbed by

a plasma discharge in the flowing propellant. The ignited plasma acts as a

lossy load, and with optimal tuning, energy absorption efficiencies over

95 percent (based on the applied microwave power) are expected. Nitrogen,

hellum, and hydrogen will be tested as propellants in the MET, at discharge

chamber pressures to I0 atm. This faclllty will be the first to test the MET

concept at these power levels, pressures, and frequency.

The performance of the laboratory MET will be enhanced with a strong solenoidal magnetic field in the throat-nozzle region, a so-called magnetic nozzle.

The magnetlc fle]d acts to compress the ionized propellant along the thruster

axis. Thls reduces wall impingement and heating and also promotes recombination of the ionic species away from the walls, with the consequent recovery of

some of their internal energy. The magnetic nozzle configuration employs a low

temperature superconducting magnet with a maximum field strength of up to 6 T,

mounted adjacent to the mlcrowave cavity. This is the first known use of a

magnetic nozzle to improve the performance of an electrothermal thruster operating at high pressure. (Applied magnetlc fields have been extensively used

to enhance the operation of low pressure magnetoplasmadynamic (MPD) thrusters.)

Calculations of the predicted performance of the MET are presented. These

calculations are made with a two-dimensional klnetlcs (TDK) computer code normally used for predictlng chemical rocket performance. The code has been

*NASA-ASEE Faculty Fellow.
extended to Include plasma reactions and recently available thermodynamic and

kinetic data for hydrogen to temperatures as high as 20 0OO K. These are the

first Known TDK calculations to be made For hydrogen plasmas. The calculations

indicate that the MET may reasonably achieve a specific impulse of 2000 sec

for hydrogen propellant at a pressure of 10 atm, a temperature of 6000 K, and

a power level of 1OO kW. The results are in agreement with basic relations for

thruster performance. The specific impulse was Found to increase with increasing chamber temperature and to depend weakly on chamber pressure, while the

thrust was found to be directly proportional to chamber pressure and to depend

weakly on chamber temperature.

II. INTRODUCTION

The National Commission on Space emphasized (ref. I) the necessity of

developlng high performance advanced propulsion technology in order to accomplish many of the important goals envisaged for the national space program.

Electric propulsion is attractive For certain of the proposed missions due to

its very high performance.

Electrothermal thrusters are one of three classes of electrlc propulsion

devices, the others being electrostatic and electromagnetic thrusters. Electrothermal thrusters include such types as the arcjet, resistojet, and the

laser- or radiofrequency-heated thruster, as well as the microwave thruster.

A schematic diagram showing the energy input, use, and loss paths in all such

thrusters is presented in figure I. The input energy, be it deposited in an

ac or dc discharge or conveyed by thermal radiation or conduction, heats the

flowing propellant to high temperature. The hot propellant then thermodynamically expands via choked flow through a throat and nozzle to produce the

device's thrust.

Electrothermal thrusters may in general be operated at relatively high

pressures, typically to many atmospheres, and thus can produce relatively high

thrust and thrust density levels. In the microwave, arcjet, laser, and radiofrequency thrusters, a plasma discharge is created and maintained by the

absorption of the applied electrical or electromagnetic energy. This plasma

efficiently heats the propellant gas passing through it or along its boundary.

For maximum thruster performance, as much as possible of the energy invested

in ionizing, dissociating, and exciting the propellant in the discharge region

must be recovered as kinetic energy during the expansion of the hot propellant

in the throat and nozzle. The hot plasma species must also be substantially

prevented from impinging on the thruster's inside walls, because this is an

important wall heating mechanism and energy loss path. The processes by which

energy is lost to the walls include electron-ion recombination, atom recombination, convection, conduction, and radiation.

Means for achieving these desirable goals include operating at high pressure, employing a propellant which has fast kinetics for its recombination and

de-excitation reactions, constricting the plasma away from the walls, and

increasing the concentration of the ionic species along the thruster axis. A

promising concept for accomplishing the latter two of these effects is the magnetic nozzle. It is broadly applicable to all types of thrusters producing

significant levels of propellant ionization.
Section III of this paper describes specific aspects of METoperation,

wlth particular consideration of operation at high power. It also presents a

description of the mechanismand expected effects of a magnetlc nozzle, as

integrated with the MET. Section IV presents details of the experimental apparatus under assembly at the NASALewis Research Center to test the METat high

power levels and with a magnetic nozzle.

In a thermal rocket, the thrust is provided by the choked-flow expanslon

of the heated propellant or combustlon gases through a throat and nozzle. The

thrust ideally varies with the chamber pressure as

F = CfAoP , (I)

where Cf is a thrust coefficient, which depends on the propellant gas speclfic heat ratio and the nozzle expansion ratio, Ao is the minimum throat

cross-sectional area, and P is the chamber pressure. This relationship presumes expansion into a vacuum. If Cf remains constant, the thrust should be

proportional to the chamber pressure.

The propellant exhaust velocity, vD, in a thermal rocket is provided by

the thermal energy of the heated propellant before expansion. Hence, from

kinetic theory, the maximum value of Vp is given by

jC3KT,_I12

Vp : \m--p-) '

(2)

where mp Is the propellant mass per expelled atom, ion, or molecule, k is

the Boltzmann constant, and T is the maximum absolute temperature of the

expelled species prior to expansion. If only a single propellant species is

present, the speclflc impulse (Isp) will thus be given by

(3kT/mp)

Isp - g

112

(3)

where g is the acceleration due to gravity, showing that the Isp should be

proportional to TI/L in thls ideal case.

In real thrusters, multiple species are usually present in the expanding

propellant. These are all in dynamic, pressure- and temperature-sensitive

equilibria with each other and the walls, or only in partlal equilibria. In

such a case, the calculatlons of thrust, exhaust velocity, and specific impulse

are much more complicated. To enable such calculations, a standardized computer code which predicts and models thermal rocket engine performance has been

developed (ref. 2). This two-dimensional klnetlcs (TDK) code has been used to

make the calculations presented in section V, for a wide range of chamber pressures and temperatures with hydrogen as the propellant. These calculations are

applicable to the microwave electrothermal thruster.
III. MICROWAVE ELECTROTHERMAL THRUSTER (MET) AND MAGNETIC NOZZLE

A. General Characteristics of the MET

The MET is an electrothermal thruster which employs CW microwave energy

to create and maintain a plasma discharge in a flowing propellant so as to

heat the propellant. As demonstrated by Arata (ref. 3), such a plasma can be

sustained at relative high pressures (>I atm) and can efficiently heat a propellant gas to temperatures greater than 5000 K, making the concept attractive

for propulsion applications. The thruster relies on the choked-flow thermodynamic expansion of the heated propellant through a throat and nozzle to produce thrust. Hence, it possesses the capabilities of high thrust and thrust

density (thrust per unit cross-sectional area of the nozzle) common to thermal

thrusters operating at high pressure. The major performance (i.e., specific

impulse) constraint of such thrusters is throat and nozzle wall material temperature llmltations.

The MET possess several potential advantages over other types of electric

thrusters. It is electrodeless; hence the erosion and energy losses associated with electric thrusters employing electrodes may be substantially eliminated. The plasma volume can, within limits, be externally controlled in size

and location by such variables as the pressure, applied power, gas flow characteristics, and applied magnetic field. Such control of the plasma allows

for optimization of the thruster performance which is not available with other

types of electric thrusters. The MET can also be operated on a wide variety

of propellants.

Very high efficiencies (>95 percent) have been demonstrated by Whitehair

(ref. 4) in absorbing the input microwave energy and converting it to propellant kinetic energy In the MET. The input energy is first absorbed by the

free electrons in the plasma discharge, and due to the high pressure environment, the electron energy quickly equilibrates with all the propellant species

present. Unless the equilibrium plasma temperature becomes high enough to

cause significant thermal ionization, the ionization remains quite low - typically less than 1 percent (ref. 5).

B. Applicators

Experimentally, MET evaluation to date has been conducted at the industrial heating frequency of 2450 MHz, with narrow bandwidth power sources. The

investigations of Arata <refs. 3 and 6), conducted at 915 MHz in nitrogen and

hydrogen, did not employ apparatus intended for thruster evaluation and did

not investigate choked flow of the working gas. Two types of microwave

applicators, the coaxial and cavity applicators shown in figures 2(a) and (b),

respectively, have been successfully employed to achieve efficient energy

absorption in the plasma. The coaxial applicator is operated in a transverse

electromagnetlc (TEM) mode and the plasma discharge acts as an extension of

the center conductor. The cavity applicator is operated in a lower order

transverse magnetic (TM) mode and the plasma discharge acts as a lossy load.

With optimal tuning, efficient coupling of the microwave power into the plasma

using both of these applicators has been achieved in nitrogen and helium at

low power levels and moderate (I to 3 arm) pressures (refs. 4 and 7).
Whenmicrowave radiation is coupled Into a resonant cavity, electromagnetic fields are established in standing wave modes. Energy is absorbed by

the plasma through acceleration of the free plasma electrons by the ac electric fields. Thus, the plasma is formed in the region where the electric

fields are the strongest. The electrons impart their kinetic energy to the

ionic and neutral species in the plasma through collisions. The higher the

pressure, the greater the collision frequency and the closer the plasma

approaches thermodynamicequilibrium. The cooler propellant gas flowing into

and along the periphery of the discharge volume is then heated by convection,

conduction, and radiation. These energy transfer processes are generically

sketched In figure I.

The electromagnetic field lines of the TMoI1 and TMoI2 modesare depicted

in figures 3(a) and (b). (The subscripts denoting the modesrelate to the

numberof half-wave variations in the azimuthal, radial, and axial components

of the electromagnetic field, respectively.) The lower order TMmodesof the

cavlty applicator lend themselves most readily to METoperation because they

concentrate the most intense electric flelds along the cavity axis, where the

plasma discharge Is formed. Thus, the plasma is located away from the dlscharge tube wall and energy losses to the walls are minimized. Stabllization

of the plasma volume at a fixed location with a fixed shape is principally the

result of the concentration of the electric field on the axis of the cavity

and the peripheral cooling of the plasma by the flowing propellant gas and any

physical walls present.

Both the TMoIl and TMoI2 modeshave been successfully utilized in cavity

applicators (ref. 4) to maintaln high-pressure discharges and both will be

investigated In the present research effort. The cavlty is designed and can

be tuned so that either the TMoI1 or TMoI2 modeIs singly excited, without

admixture of other modes. By adjusting the position of a cavity sliding short

and the insertion of the antenna coupling probe, the microwave circuit may be

exactly tuned to match the complex impedanceof the load and minimize the

reflected power from it (ref. 8). The TMoI1 modeis preferable in realizing

the METconcept. Not only is the cavity required half as long as that for the

TMoI2 mode<which will reduce wall losses), but more importantly the discharge

resides at one end of the cavity due to the electric field node at the cavity

midpoint. This allows the throat-nozzle assembly to be located outside the

cavity, yet still be in close proximity to the discharge which is heating the

propellant gas. In this regard, the coaxial applicator (fig. 2(a)) has an

advantage over the cavity applicator since the discharge may be placed arbltrarily close to the throat-nozzle assembly.

The MET concept is technically feaslble at either the 915 or the 2450 MHz

industrial heating frequencies available. In choosing between them to extend

investigations to higher power levels than previously attempted, several factors were considered. Some of these relate to the different wavelengths of

the microwave radiatlon at the two frequencies: 33 cm at 915 MHz and 12 cm at

2450 MHz. Cavity, waveguide, and coaxial transmission line dimensions all

scale with wavelength. As a result, these components are somewhat heavy and

bulky for the lower frequency. However, the larger dimensions allow for

larger discharge tubes and throat-nozzle assemblies. This factor may offer

both operational and fabrication advantages. Another important consideration

Is the cost and efficiency of high-power, low ripple magnetron generators at

the two frequencies. On both counts, the lower frequency has a distinct
edge. These considerations led to the decision to assemble the apparatus for

the present investigation to operate at the 915 MHz frequency.

C. Discharge Characteristics and High-Power Operation

Plasma discharges typically have electron number densities greater than

1012 cm-3. At such densities, their cutoff frequency (i.e., plasma frequency)

for electromagnetlc wave propagation is well above the 915 and 2450 MHz microwave frequencies. The plasma frequency electron number densities at 915 and

2450 MHz are 1.0xlO I0 and 7.5xi0 I0 cm-3, respectively. Hence, microwave radlation at either frequency will not propagate through a plasma of even moderate

electron density and will instead be mostly reflected at the plasma boundary.

Only a small fraction of the energy contained in an electromagnetic wave at

these frequencies will be absorbed by the plasma. However, it is the electric

fields of the applicators described in the previous section and not the incident electromagnetic waves that transfer the energy to the plasma. This is

considered further below.

The quality factor (Q) of a MET resonant cavity is the ratio of the energy

stored In the electromagnetic fields to the total energy losses. These losses

are to the cavity walls, discharge tube, and plasma. With optimal tuning, the

Q approaches 15 000 for an empty cavity and is reduced to less than 100 for a

cavity with a plasma load (ref. 9). Hence, the plasma acts as a lossy load

and efficiently absorbs the incident microwave power, with the energy losses

to the discharge tube and cavity walls being of the order of one percent. As

noted above, adjustment of the slidlng short positlon and the insertion of the

input power probe provide the capability of exactly tuning the impedance of the

plasma load for minimum mismatch and reflected power. Optimal tuning further

requires a narrow bandwldth, low ripple microwave signal, which present-day

magnetron generators can provide. Ripple and any other oscillating condition

causing plasma fluctuations will give rise to an ac component of the plasma

load impedance which cannot be accommodated by the dc tuning mechanisms

available.

The principal experimental varlables affecting and allowing some control

of the plasma volume in a MET operating at a fixed frequency are the discharge

chamber pressure and the absorbed CN power level. Increasing the pressure compresses the plasma volume, due to increased collision frequencies. Increasing

the absorbed power increases the plasma volume, due to increased plasma temperatures and densities. Thus, the plasma volume may be kept roughly constant

while increasing the power level by correspondingly increasing the pressure.

This is important in achieving high-power operation and permits limited external control of the absorbed power density of the p]asma discharge.

Ignitlng the MET discharge must normally be done at pressures less than

I00 torr, to achieve breakdown in the propellant gas. This in turn requires

that only modest starting power levels be applied. Otherwise, the discharge,

once ignited, will immediately expand to contact the discharge tube walls

before the pressure can be raised to prevent this. Such an occurrence could

cause melting of the walls and catastrophic failure. Nith the discharge

started at low pressure and power, both variables are simultaneously and

smoothly Increased to achieve high-power operation. During this process, the

cavity must be continually retuned to match the changing load impedance and

maintain minimum reflected power operation. At high applied power levels,
even low reflection coefficients meanthat substantial and potentially hazardous quantitles of power are being reflected back into and absorbed by the

microwave c|rcuit.

As the electron density of the plasma is increased, the penetration depth

of the electrlc flelds into the plasma Is reduced. This is knownas the skin

effect. At somehigh level of power density, the skin effect is expected to

prevent further Increase in the power density. At this point, the plasma density is so h|gh that the microwave power is absorbed in a very thin skin layer

and the plasma k|netic pressure is so high that increasing the propellant gas

pressure has no further effect in decreasing the plasma volume. Therefore, a

further Increase of the absorbed microwave power only increases the plasma

volume.

The skln effect domination of the MET discharge characteristics at high

power may provlde a benefit which more than compensates any limits on the

plasma power density It imposes. This is that it forces all energy absorption

to take place at the plasma boundary, precisely where the desired colllslonal

energy transfer to the cool propellant gas takes place. At the same time, the

skin effect protects the hot interior of the plasma from further energy absorption and heating. The result should be a relatively flat temperature profile

across the plasma, which will minimize the radiation losses which are expected

to constrain high-power operation. Hence, relatively high average plasma temperatures should be achievable for given limiting radiatlon losses.

During hlgh-power operation, the stability of the MET discharge is of

critical concern. Should plasma fluctuations or externally induced perturbatlons (such as flow instabilities or a microwave generator upset) cause it to

extinguish, it may not automatically reignite under the hlgh-pressure conditions prevailing. The cavity will then revert to a mistuned high-Q system.

At a high applied power level, discharges may be induced outside the discharge

tube or in components of the microwave circuit. Potentially disastrous heating

of the cavity walls, discharge tube, or other components could rapidly occur.

It is imperative, therefore, that a reflected power or impedance mismatch sensor circuit immediately and automatically shut off the applied microwave signal

in the event of a discharge extinction under high power conditions.

D. Magnetic Nozzle

The MET concept lends itself to potentially significant enhancement if

implemented with a magnetic nozzle. Figure 4 presents a schematic of such an

integrated system, based on a cavity applicator. A strong dc solenoidal magnetic field is configured coaxially with and adjacent to the mlcrowave cavity,

with the throat-nozzle assembly situated in the maximum magnetic field region.

A magnetic pressure directed in the opposite direction to the magnetic field

gradient acts on the charged particles in the plasma and has a magnitude of

B2/2Wo, where B is the magnetic field strength and _o is the permeability

of free space. The magnetic pressure is thus about 4 arm at 1 T (104 G) and

about 140 arm at 6 T (6xlO 4 G). If it were possible to heat a diatomic gas at

I arm to I0 000 K, then fully dissociate and ionize it, the resulting kinetic

pressure would be 73 atm. Thus, the magnetic pressure inside present-day

superconducting magnets can far exceed the gas kinetic pressure. This illustrates the effective pinching action the dc magnetic field can have on the MET
plasma discharge, an effect which scales as the square of the field strength

and is llmited only by the available field strength of the magnet employed.

The pinch effect of the magnetic nozzle can serve several useful functions in the MET. By constricting the plasma volume, it may substantially

increase the above-described discharge power density limit imposed by the skin

effect. It will also directly serve to reduce plasma-wall interactions by compressing the plasma away from the discharge tube walls. (The cyclotron radii

of I0 000 K electrons and protons in a 6 T magnetic field are 6.0xlO-5 and

2.7xl0 -3 cm, respectively.) This effect may substantially ameliorate perhaps

the severest limitation to electrothermal thruster performance: that the wall

materials cannot survive the propellant temperatures necessary for the desired

performance levels. Another benefit of the pinch effect created by the magnetic nozzle is in increasing the recovery of the ionization, dissociation,

and excitation energy expended in the discharge volume. The excitation energy

resides in the quantumstates of the species present. By constricting the

charged species along the axis of the thruster, the densities of these species

are raised there and the equilibrium and recombination reactions which release

the contained energy of these species are thereby promoted. Thus, chemical and

electrical energy which would otherwise be lost is converted into the desired

kinetic energy.

A final desirable effect of the magnetic nozzle lies in its stabilizing

the plasma discharge location on the magnetic field axis. Should the discharge volume asymmetrically expand or moveoff this axis, due to convective

or buoyancy forces for example, the axially symmetric magnetic pressure, which

is directed inward, will tend to recenter the discharge volume on the axis and

restore its axlally symmetric shape. Becauseof these expected benefits, the

apparatus built to test the thruster was designed to incorporate a superconducting magnet forming such a nozzle.

IV. EXPERIMENTALPPARATUSANDOPERATION

A. Microwave Circuit and PowerGenerator

The significant componentsof the microwave circuit to be employed for the

METlaboratory testing are shown in figure 5. The circuit is designed to operate at a frequency of g15=5 MHzand at an impedanceof 50 _. The waveguide

section of the system is constructed of standard NR975waveguide components.

The microwave generator employed is a modified commercial magnetron type

with a maximumCNpower output of 30 kN. Its output signal has a typical bandwidth of 0.2 MHz. The output power level is adjustable over three ranges

(depending on the plate transformer taps connected), providing stable operation

from about 2 to 7 kWon the low range, 5 to 12 kWon the middle range, and 7 to

30 kWon the high range. As is characteristic of magnetron generators, the

output signal center frequency decreases a few megahertz as the output power is

reduced to low levels. No significant sideband radiation is produced by the

generator. A 3-port circulator inside the generator protects the tube from

injurious back-reflected power. (As seen in fig. 5, a second, external 3-port

circulator further protects it.)

Several modifications of the generator electronics circuitry from that of

the commercial heating precursor have been implemented to achieve the required
l percent peak-to-peak ripple specification for the output signal. The principal change is the use of a dual secondary, 3-phase plate transformer with a

12-phase, full-wave rectified output. Other modifications madeto reduce the

ripple include conversion of the filament heater circuit from ac to dc and the

incorporation of substantial additional filtering in both the anode current

control and filament heater circults. In addition, careful balancing of the

phase voltages of the input 3-phase, 440 VACpower aids in reducing the rlpple.

Although the microwave power applied to the METcavity can be varied by

the generator control, this does not allow the power to be continuously

adjusted from a starting level less than I kN to an operating level near the

30 kN generator limit. In addition, the few MHzfrequency changeof the output signal from low to high power complicates maintaining the system tuned for

minimumreflected power and impedancemismatch while the magnetron output

power is increased. Furthermore, at the low end of each of the power ranges,

the output signal is less stable than at the high end.

To overcome these difficulties in continuously adjusting the power

applied to the cavity over the full generator range, an external 3-port ferrite circulator and a hybrid phase shifter-tuner are incorporated in the microwave circuit, as seen in figure 5. The 3-port circulator transmits all of the

microwave generator output power to the phase shifter-tuner port, with no significant amount of the generator power passing directly to the cavity. The

phase shlfter-tuner contains double-bucket shorting assemblies. It has one

adjustment crank to effect any desired phase shift over a 190° range and

another to tune out an Impedancemismatch in the system of up to a voltage

standlng wave ratio of 1.10. By meansof these controls, 0 to 98 percent of

the input power to the unit can be reflected back into the 3-port clrculator

and thence to the cavity. The output port of the hybrid phase shifter-tuner

leads to a water load with a greater than 30 kN CWpower absorption capaclty.

All of the input power from the generator to the phase shifter-tuner which is

not reflected back into the 3-port circulator is absorbed in the water load.

The waveguide-to-coax transition indicated in figure 5 makesa transition

from the WR975waveguide used in the generator and power adjustment section of

the microwave circuit to the flexible 12.7 cm (5 in.) coaxial cable used in the

remainder of the circuit up to the METcavity. The transition is specially

designed to prevent or minimize moding in the standard 15.6 cm <6-I/8 in.)

Electronic Industries Association (EIA) flanged connection to the coax cable.

The cutoff frequency in the connector is potentially as low as 876 MHz, somewhat below the 915 MHz operating frequency. Should a small amplitude higher

order mode be excited in the coax connector, it will probably damp out completely in the 4.6 m length of the coax cable leading to the MET cavity. However, it remains an undesirable power loss and source of heating in the coax

line and connector.

The 12.7 cm flexible coax line has an air dielectric and is rated for

approximately 32 kW of CW power at 915 MHz under laboratory conditions. It

has a cutoff frequency of about 960 MHz and over its 4.6 m length, it attenuates the 915 MHz signal by less than 0.05 dB.

In conducting an experiment with the MET, it is anticipated that the generator output power level will be set at slightly more that the maximum

thruster operating level required, with the phase shlfter-tuner set for 0 percent reflection of this power. All of the generator output power will thus be
dumpedInto the water load (where the power level will be measuredaccurately

by calorimetry). The phase shifter-tuner will then be adjusted to reflect the

proper low level of power to the cavity for starting the METdischarge. After

the discharge has been started and stabilized, the phase shifter-tuner will

be further adjusted to raise the level of power applied to the cavity semicontinuously up to the desired operating level. Whenthis point is reached,

only a small fraction of the generator output power will then be absorbed by

the water load.

After the METdischarge has been started and frequently while the power

to it is being raised, the microwave circuit will be retuned for minimumimpedance mismatch and reflected power. Both the tuning adjustments on the cavity

and that of the phase shifter-tuner will be used to accomplish this. As previously noted, adjustments of the chamberpressure will also normally be

required during the power increase. Noadjustment of the output power level

of the microwave generator during an experiment is expected to be needed.

B. Microwave Cavity and Discharge Tube

The microwave cavity and discharge tube, as integrated with the superconducting magnet for the magnetic nozzle investigation, are depicted in figure 6.

The cavity is a cylinder 45.7 cm in inside diameter by 57.2 cm long. It has a

screened viewport located to facilitate visual or spectroscopic observation of

the discharge region. The sliding short at the back of tile cavity moveson

rollers and is driven by a drive and pinion gear arrangement with three lead

screws. This mechanismmaintains the slidlng short always parallel to the

fixed front plate of the cavity and prevents it from binding while being moved.

The drive gear is manually driven via another gear, enabling fine positioning

control and reproducibility with minimal backlash.

The sliding short has a total travel of 24.1 cm, allowing the cavity interior length to be adjusted from 16.5 to 40.6 cm. The calculated approximate

shorting locations to tune for the TMoII and TMoI2 modesto be investigated are

at interior cavity lengths of 19.6 cm and 39.2 cm, respectively. Electrical

contact between the shorting flange circumference and the cavity cylindrical

wall is maintained during flange translation by beryllium copper fingerstock.

The microwave power input port of the cavity is centered 9.8 cm from the

front plate, at the expected location of the downstreamelectric field maximum

in the TMoII mode. The 12.7 cm coaxial cable carrying the microwave signal is

directly connected to the port via a 15.6 cm (6-I/8 in.) EIA flanged connection. The connectors are modified and the port is designed to maintain the

samecoaxial cable average dimensions for the outer conductor inner diameter

and the inner conductor outer diameter through the connection and port into

the cavity itself. Again, this is to preclude the excitation of unwanted

higher modesin the unmodified connectors, modeswhich could then propagate

into the cavity and interfere with the clean excitation of the single desired

cavity mode.

An extension on the Input port center conductor, having a hemispherical

end, extends into the cavity and acts as an antenna for the input signal. A

captivated threaded ring joining the coupling port and the connector portions

of the assembly enables fine tuning adjustment of the excitation probe insertlon into the cavity. This insertion maybe varied from 2.5 to 5.1 cm,

lO
measuredfrom the cavity inner circumference. Fingerstock attached to the

Inner circumference of the port outer conductor maintains the proper dimensions

and electrical continuity of the outer conductor across the sliding joint at

the tuning rlng.

The microwave cavity cylindrical body, front flange, sliding short, and

coupling port are all fabricated of oxygen-free high conductivity (OFHC)copper. All the interior surfaces of these componentsare finish-machined and

polished to a 0.4 pmfinish and then coated with a I00 pmthick layer of silver. Th|s surface treatment is intended not only to assure optimal electrical

performance of the cavity but also to makethe interior surface highly reflective to the considerable thermal radiation it receives from the discharge tube.

All componentsof the cavity are water-cooled on their exterior, enabling the

cavity to absorb the full 30 kN output power of the microwave generator without

overheatlng, in the event of a malfunction in the cavlty operation.

The discharge tubes employed in the METand seen in figure 6 must be fabricated of materlals having a high melting point, high strength, and low

dielectric loss tangent. The two-tube system shownin the figure extends completely through the cavity and is concentrically positioned on the cavity axis.

The outer tube acts as a vacuumcontainment wall and is fabricated of fused

quartz so as to be optically transparent. The inner tube Is the actual discharge containment structure. For initlal low power testing, it will also be

fabricated of fused quartz, enabling spectroscopic and optical observation of

the discharge through the cavity viewing port.

Subsequent high-power testing will require an inner discharge tube fabricated of a higher temperature material meetlng the other requirements as well.

Two good candidates are fused alumina and boron nitrlde, both of which can

withstand working temperatures up to about 2000 °C in nonreactive environments.

These are, however, opaque materials, so the discharge produced in these tubes

cannot be directly observed through the cavity viewing port. The inner discharge tube initially tested will be 4.7 cm in inner diameter. This dimension

may be increased or decreased for subsequent testing, depending on the dimensions and characteristics of the plasma discharge created inside.

The vacuum jacket between the inner and outer discharge tubes will nearly

eliminate convective or conductive heat loss radially from the inner tube.

Radiative heat loss from it will also be minimized by the transparency of the

outer tube and the high reflectivity of the interior surfaces of the cavity.

Hence, the inner tube is expected to become quite hot during testing at elevated power levels. This is desirable from the point of view of achieving the

best possible thruster performance, because it implies the lowest possible

thermal losses. However, with the ultimate goal of testlng to discharge pressures as high as lO atm at power levels up to 30 kW, the material properties

of the inner discharge tube will be severely tested, and new and improved tube

materlals may have to be found.

C. Throat-Nozzle Structure and Magnetic Nozzle

The throat-nozzle structure of the MET laboratory apparatus is attached

directly to the end of the inner discharge tube, as seen in figure 6, and is

located just outside the microwave cavity front wall. A high-temperature

mechanical seal mates the two units together. The throat-nozzle structure is

II
of one-piece construction, is relatively massive, and will be fabricated of a

refractory metal such as tungsten (thoriated), rhenium, or molybdenum. Figure 7 shows the proposed cross-sectional geometry of the structure. The

geometry of the throat region is the same as that used for the performance calculations discussed in the next section, including the 1.0 mm throat diameter.

The throat-nozzle structure, as well as the whole front half of the inner

discharge tube to which it is attached, will be suspended on the common axis

of the microwave cavity and the superconducting magnet by a refractory metal

wire fastened to the top of the magnet bore liner. This minimizes the conductlve heat loss through the suspension mechanism and allows the assembly to run

as hot as possible to achieve maximal thruster performance. The massiveness

of the throat-nozzle structure is intended to bring it to a stable equilibrium

temperature during steady-state MET testing, a temperature which is insensitive to minor perturbations in the operating parameters. The changes in the

structure's steady-state temperature due to changes deliberately made in the

experimental conditions are important results to be obtained in the testing.

Such temperature changes, for example, directly indicate the magnetic nozzle

effect in reducing plasma-wall interaction heating.

The magnetic nozzle is implemented in the experimental apparatus with a

superconducting magnet, as shown in figure 6. The magnet is a short solenoid

with a clear through-bore and is mounted directly against the front flange of

the microwave cavity, with the magnet axis congruent with the cavity and discharge tube axes. Figure 8 is a photograph of the magnet structure with the

room temperature bore liner removed, exposing the liquid nitrogen-cooled inner

liner and shield. Figure 9 presents a horizontal cross section of the magnet

at its midplane. Shown are the liquid hellum-cooled coils and shield, the llquld nitrogen-cooled shield, and the room temperature outer shield and flanges.

Vacuum jackets separate and insulate all the components at different temperatures. Not shown in figure 9, but Indicated in figure 6, is the water-cooled

magnet bore liner which completes the room temperature shield through the magnet. Its inner diameter is 7.01 cm and the length of its cylindrical section

is II.6 cm. The bore liner, made of OFHC copper, has a greater than 25 kN

cooling capacity and a highly polished interior surface, designed to reflect

the maximum possible fraction of the radiant heat reaching it (mainly from the

throat-nozzle structure).

The magnetic field map presented in figure 9 shows the strength of the

axial field component during maximum field operation. The isostrength contours were derived from detailed field component values calculated for the

experimental magnet by a code which takes into account the current flow in the

windings as actually constructed. The calculated values have been verified by

field measurements. The maximum center line field strength is 5.7 T, though

field strengths up to 8 T are found in the NbTi magnet windings. The inductance of the magnet is 7.9 H and the current in the windings at maximum field

is 117.3 A.

The magnet power supply enables charging the magnet either under voltage

control, at a maximum of 5 V, or at a constant current ramping rate, up to a

maximum of 15 A/min. Full charging under the normal current ramp control

therefore requires slightly less than 8 min. The power supply also controls

the discharglng of the magnet, in the same manner and at the same rates as in

charging it, by absorbing the magnet's stored energy under controlled conditions. The current flow in the magnet, and hence the magnetic field polarity,

12
may be automatically reversed. This entails ramping the current down to zero,

swltching polarity, and ramping the current back up. At full fleld conditions, slightly less than 16 min, is required to accompllsh this. The magnet

displays negligible hysteresis, so the field strength is accurately indicated

by the current in the windings, whatever the history of the current or field

variation.

At any current or field strength up to the maximum,the magnet maybe

placed in persistent modeoperation. This is accomplished by turning off

heater power to a superconducting short between the magnet leads, located in

the liquid helium-cooled magnet chamber. The current in the magnet wlndlngs

is thus isolated from that in the power supply circuit, and the latter maybe

rampeddownto zero and the leads to the magnet disconnected, if desired. In

persistent mode, the magnet current and field strength show no measurable

changeor decay wlth time. The liquid hellum and liquid nitrogen consumption

rates !n persistent modeoperation at full field are approximately 170 and

400 cm3/hr, respectively. The magnet is protected in the event of a quench,

during which part or all of the windings suddenly becomenormal, by the

current-carrying capability of the copper matrix in which the NbTi filaments

are imbeddedwhenformed as wire for the windings.

The point at which the throat-nozzle structure is suspendedfrom the top

of the magnet bore liner is adjustable. As seen from figures 6 and 9, this

allows the throat to be placed at the maximumfleld location in the magnet or,

alternatively, either in the converging field upstream of this location or in

the diverging field downstreamof it. The effects of all these placements on

the METoperation - in particular, on the heating of the throat-nozzle structure - will be investigated.

D. Other Experimental Detalls

The propellant gas feed system for the METlaboratory apparatus is a dualline system with independent control of the gas flow in each line. One line

supplies an axially directed inflow to the discharge tube and the other provides a swirling inflow to it. The gas flow rates in each line are controlled

by massflow controllers which are insensitive to the discharge tube pressure.

Nitrogen, helium, and hydrogen, as pure gases, will be investigated in the initial experiments with the METapparatus, at flow rates up to 25 000 sccm in

each channel. The feed system is capable of supplying these flow rates with

the discharge tube pressure as high as I0 atm, but initially, this pressure

will be controlled at 3 arm, or lower. The gas flow rates, along with the

heating of the gas in the discharge tube and throat-nozzle structure, control

the discharge tube pressure, because a choked-flow condition is maintained in

the throat. In later experiments, two different gases may be run through the

gas feed system lines and mixed upon flowing into the discharge tube.

After expanding out through the throat-nozzle structure, the hot propellant gas from the METapparatus passes through a 30 cmgate valve (shown in

fig. 6) and into a large vacuumtank. Figure 10 shows the general configuration of the experimental apparatus and the tank. The tank is approximately

5.8 m long by 3.0 m in diameter. The hot exhaust gas impinges on a liquid

nitrogen-cooled cryoliner covering the opposite wall of the tank. The cryoliner acts as a heat exchanger and has adequate liquid nitrogen capacity to

13
accommodatethe full 30 kWpower output of the microwave generator. This

requires about 5.4 I/min of liquid nitrogen flow.

The exhausted propellant gas, after being cooled by the cryoliner, is then

pumpedby the vacuumtank's oii diffusion pumps, blower, and roughing pumps.

The six diffusion pumpshave a maximumtotal pumping speed of 300 000 ]/sec.

The blower pumping speed is 2350 I/sec. The tank's pumpingperformance has

been tested with room temperature helium and nitrogen, both with and without

the diffusion pum_soperating. Nith the diffusion pumpson, the limiting tank

pressure of 5x10-4 torr was reached at flow rates of 7000 sccmfor he]ium and

4000 sccm for nitrogen. With the diffusion pumpsoff, the limiting tank pressure of 0.5 torr was reached at flow rates of 200 000 sccmfor he]ium and

I00 000 sccm for nitrogen.

The 0.5 torr tank pressure upper limit does not significantly restrict

the METtest operation due to back pressure effects. The discharge tube pressure under such conditions w111 be greater than 1 arm, or more than I000 times

higher than the back pressure. Hence, the choked flow of the propellant

through the throat-nozzle structure should be unaffected. Such back pressures

may, however, increase the conductive and convective heat transfer from the

inner to the outer discharge tube across the vacuumgap between them. The

back pressure mayalso cause someseparation of the exhaust at the downstream

end of the nozzle, due to the high exit area ratio (see fig. 7).

Three meanshave been implemented to stabilize the experimental METoperation and minimize discharge plasma fluctuations. First, the ripple on the

input microwave signal has been reduced to a very low level. Second, the

applied dc magnetic field from the superconducting magnet will strongly act on

and constrain the plasma. Finally, the swirling inflow componentof the propellant gas will tend to establish axial symmetryof the gas flow through the

dlscharge tube and maintain a cool gas layer at the wall. These latter two

factors should be especially important in overcoming convective and buoyancy

effects which have previously been observed (ref. 8) to disturb the axial symmetry of the plasma volume and make it unstable in a horizontal laboratory

configuration.

Prior to starting the discharge in the METapparatus for a typical experiment, swirling inflow of the propellant gas will be introduced and the magnet

will be charged to a field strength at least sufficient to stabilize the discharge, when started. The discharge will be ignited and the applied power,

propellant flow rate, and discharge tube pressure raised to the desired test

levels as previous]y described. During this process, the microwave cavity and

circuit will be continually retuned to minimize reflected power.

Whenthe desired test conditions are reached, the magnetic field, the

total propellant flow rate, and the split of the gas _nflow betweenaxial and

swirl componentsmay be varied. Actual test conditions of propellant flow

rate, discharge tube pressure, and applied power wili be defined on the basis

of the performance calculations described in the following section. Nhenan

experiment is concluded and the microwave power turned off, the propellant

flow will typically be left unchangeduntil the discharge tube and throatnozzle structure have cooled downto room temperature, at which point a cold

flow measurementof the discharge tube pressure will be taken. The hot and

cold pressure measurementsat constant flow allow calculations to be madeof

14
the ideal thrust, specific impulse, and energy efficiency for the experimental

conditions (ref. 8).

Simple diagnostics will initially be employed to characterize the MET

operating conditions and performance. The microwave generator output power

will be measuredcalorimetrically in the water load attached to the hybrid

phase shifter-tuner while no power is applied to the cavity. The power dumped

Into the water load will also be frequently measuredwhile the discharge is in

operation. Wlth the reflected power knownfrom reading the sensor connected

to the reflected power port of the in-line dual directional coupler (see

fig. 5), accurate measurementsof the microwave power transmitted to and

absorbed In the cavity will always be available. A less accurate measurement

of the power transmitted to the cavity is also provided by the sensor connected to the forward power port of the dual directional coupler.

All the other cooling circuits in the experimental apparatus wlll also be

instrumented for calorimetric measurements. This applies as well to the cryoliner heat exchanger in the vacuum tank. Its liquid nitrogen usage rate and

outlet temperature approximately measure the collected thruster exhaust beam

power. At the elevated power levels of the intended testing, all of the above

calorimetric measurements should be quite accurate. With sufficiently complete calorlmetric data from all the cooling circuits and sufficiently low

losses in the uncooled components of the apparatus, an energy balance may be

constructed giving an estimate of the thruster thermal efficiency and all of

its energy loss paths.

Other dlagnostics measure the discharge chamber pressure and provide the

propellant flow rates In both channels of the gas feed system. The dc magnetic field strength as a function of location is given by the magnet current,

In conjunctlon with the field map.

The equilibrlum temperatures of the throat and nozzle of the throatnozzle structure will be obtained via optical pyrometry. These measurements

should reveal and quantify any magnetic nozzle effect in reducing plasma-wall

interactions in these regions. A pyrometer will also be employed through the

cavity viewport to observe the inner discharge tube exterior surface (and the

discharge itself in the case of a fused quartz inner tube). Data from these

observations will give the temperature distribution along the inner discharge

tube in the region of the discharge. When opaque inner tubes are employed,

this data will be particularly valuable in determining the location and extent

of the plasma discharge.

In later experiments, additional and more sophisticated diagnostics will

be implemented to better characterize the performance and properties of the

thruster. These include the use of emission spectroscopy and other noninvasive diagnostics to measure the concentrations and temperatures of the various

species in the discharge region, as well as in the throat and nozzle regions.

Of equal importance, a direct thrust measurement system is planned. This will

provide experimental thrust measurements which also will give values for the

specific impulse (Isp). These experimentally determlned values should be of

much improved rellability compared wlth the calculated ideal values obtained

from measurements of the hot and cold discharge chamber pressure at constant

flow rate.

15
V. CALCULATEDPROPULSIONPERFORMANCE CHARACTERISTICS

The propulsion performance characteristics of the METthruster have been

predicted w_th hydrogen as the propellant. These calculations were madewith

a two-dimenslonal klnetics (TDK) computer code usually used for predicting

chemical rocket performance. The code was extended to include reactions

involving electrons, Ions, and third bodies. These are the first knownTDK

calculations for hydrogen plasma flow through a nozzle. Previous efforts have

been llmited to one-dimensional kinetics (ODK) calculations (ref. IO). The

code ut_lizes recently calculated high-temperature (up to 20 0OOK) thermodynamlc data for hydrogen (ref. II). It also computes performance characteristlcs for one-dlmensional frozen flow (ODF), one-dimensional equillbrium (ODE),

and one-dimensional kinetics (ODK)analyses. The reaction rates used in the

kinetics calculatlons are given in table I (ref. IO). Results of these computations are appl_cable to all electrothermal propulsion concepts utilizing

hydrogen propellant.

The initial condltlons for all the calculations assumethermodynamic

equllibrium In the chamberregion for a given temperature or enthalpy at an

asslgned pressure. The equilibrium conditions are calculated by the chemical

equ111brium code (CEC) subprogram. The ODFcalculations then are madeassuming that the thermodynamic equilibrium concentrations of the various species

In the chamberare maintained during expansion through the nozzle (i.e., the

specie concentrations are frozen at the chamberconditions). The ODEcalculations are madeassuming that thermodynamlcequ_librium is reestablished at

every point throughout the expansion process. The ODKcalculations use the

reaction rates to obtain better estimates of the specie concentrations during

expansion through the nozzle. The TDKcalculatlons also use the reaction

rates to obtain estimates of the specie concentrations during expansion and

compute the performance losses due to the geometry of the chamber-throatnozzle walls. The method of characteristics is employed in the TDKcalculations to compute the properties of the expanding propellant along streamlines

which span the throat-nozzle volume. The chamber-throat-nozzle assembly geometry parameters used are shown in figure 11. The actual values employed for

these parameters closely describe the throat region of the chamber-throatnozzle structure depicted in figure 7, including the 1 mminner diameter

throat. All of the results presented here are for a subsonic or contraction

(chamber-to-throat) area ratio of 5 and a supersonic or expansion (nozzleto-throat) area ratio of 75.

The ODEresults give the upper bound on the performance, because the maximumamount of excitation, dissociation, and ionization energy in the propellant

species is recovered while thermodynamic equilibrium is maintained throughout

the propellant expansion and cooling in the nozzle. The ODF results, on the

other hand, give the lower bound on the performance, because none of the above

types of nonklnetic energy are recovered during expansion. The ODK results are

always intermediate between the other two and reflect the extent to which the

nonkinetic energy is actually recovered. The TDK results, which are the most

realistic, are slightly lower than the ODK results since they take into account

the losses due to the throat-nozzle wall curvature during expansion.

The vacuum specific impulse as a function of chamber temperature at a

chamber pressure of 10 arm is given in figure 12. The specific impulse shows

roughly the TI/2 dependence on temperature predicted by equation (3). The

16
Iso was found to dependweakly on the chamberpressure. The results shownin

th_ figure indicate that a specific impulse of 2000 sec can be reached for a

chambertemperature of about 6000 K. Increasing the chambertemperature from

6000 K to 12 000 K does not gain muchimprovement in the specific impulse.

This is due the fact that increasing the enthalpy, in this range, results in

increased ionization, not greater kinetic energy, of the hydrogen in the chamber, and little of the ionization energy is recovered during the expansion

process. The specific impulse from the kinetics calculations is seen in figure 12 to be near the equilibrium limit at the lower temperatures and near the

frozen flow limit at the higher temperatures, again indicating the extent to

which the ionization energy is not recovered during expansion.

The thrust as a function of chamberpressure at a chambertemperature of

4000 K is shownin figure 13. The thrust is seen to be directly proportional

to the pressure, in agreementwith equation (I), and was found to depend

weakly on the chambertemperature. The thrust and specific impulse tradeoff

at a constant power (i.e., input enthalpy) of 107 kWis shownin figure 14.

This relationship agrees with the basic propulsion relation which predicts

that specific impulse and thrust are inversely proportional to each other.

This tradeoff provides a valuable capability for electrothermal thrusters:

high thrust TeveTs may be obtained at the expense of specific impu/se by

increasing the propellant mass flow rate.

The results shown in figures 12 to 14 do not include plasma-wall interactions, which will lower the performance of the MET. Such interactions include

electron-lon and atom-atom recombination, heat conduction and convection, and

radiation transfer. The MET experimental apparatus described in the previous

sections includes several features, such as the magnetic nozzle, which are

designed to reduce these wall interactions. When the wall interactions are

sufficiently reduced, the e×perimental results are expected to approach the

TDK predicted performance for the throat-nozzle configuration tested, taking

Into account boundary layer losses.

Vl. CONCLUDING REMARKS

This paper describes the microwave electrothermal thruster (MET) concept

and apparatus being assembled at the NASA Lewis Research Center to test it.

It also presents the predicted MET performance as calculated using a thermal

rocket performance computer code. The laboratory apparatus is designed to

operate at 915 MHz and at CN power levels up to 30 kW. This type of thruster

has not previously been evaluated at frequencies below 2450 MHz or at power

levels more than about I0 percent of that intended in this investigation.

In the MET laboratory apparatus, the microwave energy is applied in a cavity applicator and is absorbed by a plasma discharge in the flowing propellant.

The TMoI 1 and TMoI 2 standing wave modes will be employed in the testing. They

concentrate the electric field on the cavity axis where the discharge is

desired. A low ripple microwave signal is required to minimize plasma fluctuations and tuning losses. The ignited plasma acts as a lossy load, and, because

of this, energy absorption efficiencies greater than 95 percent are expected.

The plasma discharge is started at low power and pressure, and then both

variables are smoothly and simultaneously raised to the operating levels. This

keeps the plasma volume controlled and roughly constant in size. During the

17
the power increase, the cavity must be continually retuned to minimize

reflected power in the system. The power applied to the cavity is principally

varied by meansof a hybrid phase shifter-tuner in the microwave circuit. The

propellant gases to be tested are nitrogen, helium, and hydrogen, at discharge

chamberpressures up to IO atm and flow rates initially up to 50 O00 sccm.

A magnetic nozzle is incorporated in the experimental apparatus and is

expected to slgnificantly improve the METperformance. The magnetic nozzle is

created by a short, solenoidal superconducting magnet generating a maximum

on-axls field strength of 5.7 T, applied in the throat-nozzle region. The

pinch effect and high magnetic pressures generated by the magnetic field have

the potential to reduce wall losses from the hot plasma species; promote

recovery of ionization, dissociation, and excitation energy from the hot species; and stabllize the plasma discharge in the cavity. The stabilizing

effect of the magnetic field on the plasma will be augmentedfluid-dynamically

by the use of a swirl componentin the propellant flow.

The skin effect, which limits microwave energy absorption to the outermost layer of a dense plasma discharge, is expected to play a major role in the

high power operatlon of the MET. This effect may limit the achievable power

densities in the plasma discharge but it also is expected to yield a uniform

temperature profile across the plasma volume, with the energy absorption taking

place at the plasma surface where the energy must be transferred to the surrounding unheated propellant. Radiation losses from the plasma should be minlmized by this mechanism. The magnetic field of the magnetic nozzle is expected

to compress the plasma volume beyond the skin effect limit, thus permitting

higher power density operation of the MET.

The propulsion performance of the METwith hydrogen propellant has been

predicted with a two-dimensional kinetics (TDK) computer code which incorporates one-dimensional equilibrium (ODE), one-dimensional frozen flow (ODF), and

one-dimensional klnetics (ODK) subprograms. The ODEresults give the upper

bound on the performance, in which the maximumamountof the excitation, dissociation, and ionization energy invested in the propellant in the discharge

chamberis recovered as it expands and cools in the nozzle. The ODFresults

give the lower bound on the performance, in which none of the above types of

nonkinetic energy is recovered during expansion. The ODKresults are intermediate betweenthe other two and reflect the extent to which the nonkinetic

energy actually is recovered. The TDKresults, which are the most realistic,

are lower than the ODKresults and take into account the losses due to the

throat-nozzle geometry during expansion.

The specific impulse was found from the TDKcalculations to increase

approximately as the square root of the chamber temperature and to depend

weakly on the chamberpressure. A specific impulse of 2000 sec is achievable

by the METwith a bulk gas chambertemperature of 6000 K, at a chamberpressure of IO arm. The thrust was found to be directly proportional to the chamber pressure and to depend weakly on the chambertemperature.

The TDKcalculations to date have not taken into account wall losses due

to recombination and thermal heating effects. The METexperimental apparatus

is designed to minimize these losses, by such meansas the magnetic nozzle.

Whenthe wall interactions are sufficiently reduced, the experimental results

are expected to approach the performance predicted by the full TDKcode.

18
REFERENCES

I. Pioneering the Space Frontier, U.S. National Commissionon Space, Bantam

Books, NewYork, 1986.

2. Nickerson, G.R., et a1.: Two-Dimensional Kinetics (TDK) Nozzle

Performance ComputerProgram, Software and Engineering Associates, Inc.,

Carson City, NV, February, 1988.

3. Arata, Y.; Miyake, S.; and Kobayashi, A.: High Power Microwave Discharge

in Atmospheric Hydrogen Gas Flow. J. Phys. Soc. Japan, vol. 44, no. 3,

Mar. 1978, pp. 998-1003.

4. Nhitehair, S.; Asmussen,J.; and Nakanlshi, S.: Microwave Electrothermal

Thruster Performance in Helium Gas. J. Propulsion Power, vol. 3, no. 2,

Mar.-Apr. 1987, pp. 136-144.

5. Chapman,R.: Energy Distribution and Transfer in Flowing Hydrogen

Microwave Plasmas. Ph.D. Thesis, Michigan State Univ., 1986.

6. Arata, Y., et al.: Research of a Stationary High Power Microwave Plasma

at Atmospheric Pressure. J. Phys. Soc. Japan, vo]. 40, no. 5, May 1976,

pp. 1456-1461.

7. Asmussen, J; and Whitehair, S.: Experiments and Analysis of a Compact

Electrothermal Thruster. IEPC Paper 88-101, Oct. 1988.

8. Nhltehair, S.: Experimental Development of a Microwave Electrothermal

Thruster. Ph.D. Thesis, Michigan State Univ., 1986.

9. Frasch, L.L.: An Experimental and Theoretical Study of a Microwave Cavity

Appllcator Loaded with Lossy Materials. Ph.D. Thesis, Michigan State

Univ., 1987.

I0. McCay, T.D.; and Dexter, C.E.:

Hydrogen Laser Thermal Thruster.

July-Aug. 1987, pp. 372-376.

Chemical Kinetic Performance Losses for a

J. Spacecr. Rockets, vol. 24, no. 4,

11. McBride, B.J.: Personal Communication, NASA Lewis Research Center,

Cleveland, OH, 1988.

TABLE I. - HYDROGEN REACTIONS AND KINETIC DATA

[M is a third body: k is the reaction rate, with units of cm 3,

K, mole, and sec; hv is a photon.]

Reaction

H+H+M > H2 +M

H+ + e- + M > H + M

H2 + e- ) H + H + eH-+H ) H2 +ell+ + e- ) H + hv

H+ + 2e- ) H + eH++H - ) H+H

Reaction rate

k = 6.40xi0 l? T -]'0

k 5.26xi026 T-2'5

k 1.91x]O }l T e (-I02"T)

k = 7.83xi0 TM

k : 3.77xI0 ^13 T-0.58

k l 08xl_ 3_ T -4"5

k I]30xI_18 T-O.5

19
EXPANDED GAS

REGENERATIVE

THROAT-NOZZLE

COOLING-.,,

VIEWING _ '

WINDOWJ'_

CENTER COND

(IMPEDANCE TUNING

ADJUSTMENTS NOT SHOWN)

/--_ THROAT-NOZZLE

ASSEMBLY

/

_.. _-MICROWAVE POWER

CONTAINMENT GUIDE

/"

"-PROPELLANT GAS

INFLOW t- (PARTIAL)

MICROWAVE

SIGNAL

INPUT

(a) Coaxial.

ABSORPTION PROPELLANT

OF APPLIED

ENERGY-_

PL_AP_EU__-- APA

ELECTROMAGNETICoR //; ['_,_

ELECTRIC,_ /Ill V/t ;-_

ENERGY= // _ _ Y,4 / I

_--'IENERGY // \/_<_-J_//

H_XCHA"GE II:/"-.'_-"

IIPROPEL"NT11 ._D

II .'11 PROPELLANT

P .-'; II //_ .." \\v ./'j

4--DISCHARGE _ f_-THROAT

OR HEATIM3_

r

CHAMBER "_/-- NOZZLE

EXPANS ION

Figure 1. - Electrothermal thruster schematic.

APPLICATOR

/

I d

PAD_TION

,-DISCHARGE

(NONE FOR

RESISTOJET

v

RADIAL HEAT

CONVECTION

AND

CONDUCTION

I

MICRO- _=R===

WAVE

SIGNAL

INPUT aI ..

PROPELLANT GAS

INFLOW

ANTENNA

)

,,'- SLIDING

SHORT

"-CAVITY

SCREENED

VIEWING s s

I PORT _s r

%.

_- THROAT/NOZZLE

EXPANDED GAS

(b) Cavity.

Figure 2. - Microwave thruster applicatc,'s.

2O
CROSS SECTION OF

DISCHARGE TUBE

FIELD _--//,_

U

(a) TM012 . (b) TMol 1.

Figure 3. - Elec_'nagnetlc field patterns for MET modes, with discharge

placernent.

r- SOLENOIDAL

MAGNET

I _ I .._,--- B FIELD

EFIELD t_, , , A ,, _ % _ r-'_.."/" (DC_

(AC)_,,L_J J / _ 11 /"_ \ I...,. ,, ,, " "

BAF_ELD'J-_a f __ CYCLOTRON::_: " :!_i!ii.:ii_'.'i!_.:i!! ...... _:':+:':':': .....

I METCAVITY I MAGNETIC NOZZLE I

(TM0_ 2 MODE)

Figure 4. - Magnetic nozzle schematic, as Integrated wl_ cavity-MET.

ALL WAVEGUIDE

WR 975

MICROWAVE GENERATOR

0-30 kW, 915_+ MHz

WAVEGUIDE TO _i i liiIIII]11111-[-IIhHrIrI'T'fI'ITT-I'FI-TIIrIIIHTHTIr_ ;'_ ^r,. neT^=.=

COAX TRANSITION I III 151_. FLEXJBLECOAX TRANSMISS=ION L_NEt/_ _)_N_E"F_I_N_I_T

_' - - '_ ANTENNA PORT

90 ° STEP '___/'/ty%

TWIST

FORWARD _ r_o--_o_ DI 7 M/CV??yWAVE

REFLEC TED _'_ / D'URALTIoNAL ( _ I

90 ° STEP c I ,Ill

REFLECTED

POWER

3-PORT CIRCULATOR

• SHORTING

[_ _ SHIFTER-TUNER _ _ I I I - BUCKET

WATERI WAT L,--II .--..--'t II I _ _ ADJUSTMENT

OUT LOA[ -I_J-U U

t

WATER

IN

Figure 5. - Microwave circuit for MET test apparatus.

21
WATER

COOLING_

MAGNETCRYOSTAT

F SHORTING FLANGE

,t (MOVABLE)

.-r TM011 LOCATION

SCREENED

VIEWPORT

= •

i

t MICROWAVE CAVITY

I

I

I

I

i

I

!

I

I

I

, WIRE

-- VACUUM LINER SUSPENSION-

( I

GAS PROPELLANT _ DISCHARGE TUBE _;_"D I'SCI'_"I_G E.X"_'X_ iNFLOW _\\\'-'-'-'-'_'-_'_'-'-'-_

(AXIAL AND SWIRL)

/-TM012 LOCATION

MICROWAVE SIGNAL

INPUT ANTENNA

I (INSERTION ADJUSTABLE

F FINGERSTOCK t (/_II

.U

II

Ii

II

It

i I

i I

II

/

GATE

VALVE

__ F MAGNET

WINDINGS

I

_ WATER-COOLED

BORELINER

l

- t

Figure 6. - Cavity-MET apparatus integrated with superconducting magnet.

lo cm

ROOM

TEMPERATURE

SHIELD ----"

r- HIGH TEMPERA'R/RE

DISCHARGE /-" MECHANICAL SEAL WIRE SUSPENSION

TU=7_" _,OT,O=_ _s.ow.

• r-DOWNSTREAM

,,

I= I 135.1 =I

_T

LOCATION

-nomde _ =eemeW.

)

LIQUID

NITROGEN

COOLED

SHIELD J

Figure 8. - Superconducting magnet.

10¢m

22 '": ......... - P_CE

L, /'.':!< ;.,"',!t , :. :,: p,
340 mm -'

/ I I ,_OOLEOSH,ELO !

/ _ / \ _LANGE /

r-_wr_71 I_ ,-VACOOM/

]_4_ _ 3o,--'--$.,'/I-_,'-,._LL/I_AC_ETS /

I 79_m H II II I _ J kllll \ VALUE /" IIIH _ / _ I _,I_ _ FILLED

U

...... -' _ 3.0 _ I-- _, ' SHIELD

ROOM TEMPERATURE BORE _ _ / \ _- FLANGE

UNER NOT SHOWN _, 20 /

_. _ Bz VALUES IN TESLAS

Figure 9. - Cross section and axial component field slrength map of superconducting magnet.

PROPELLANT GAS

INFLOW FROM

DUAL-LINE FEED SYSTEM

SHORTING FLANGE

TUNING MECHANISM _'_

n

DISCHARGE J

LHe AND LN 2

RESERVOIRS

VACUUM TANK

(30 mid x 58 m long)

_MAGNET BORE LINER

_THROAT-NOZZLE STRUCTURE

LN2 -COOLE[

CRYOLINER

MICROWAVE SIGNAL f"

INPUT ANTENNA /

/

TUNING

MAGNET

FLEXIBLE

COAXLINE

GATE

VALVE

F4 SIX DIFFUSION PUMPS

Figure 10. - Microwave electrothermal thruster test apparatus and vacuum facility.

NOT TO SCALE

23
I

6i e

el = 25 ° THROAT RADIUS (rt) - I mm

8 - 15=

rI = 2 mm CHAMBER AREA RATIO (rc/rt) 2.5

rd - 4 mm

ru - 2 mm EXPANSION AREA RATIO (r/rt) 2.75

Rgure 11. - Chamber-throat-nozzle usembly geomeVy for "FDK¢alcukltlons.

6OOO

r rd _ 5000

r--_ - 3000

o

2OOO

1000

CHAMBER PRESSURE = 10 atm

_

_ o_[] TDK

I I I I

,5000 10,000 15,000 20,000

CHAMBER TEMPERATURE, °K

Figure 12. - Spedflc Impulse as a fur)c_on of chamber ternpem_ure.

CHAMBE R TEM

z

l _ A OOE

2_- _ oo_

0 5 10 15

CHAMBER PRESSURE, aim

FLgure 13. - Thrust as a function of chamber pressure.

.=

6000

500O

40OO

3000

2O00

IOO0

POWER. 107 kW

OOOF

Z_OD6

OOOK

rl TDK

I I I I

5 10 15 20 25 30

THRUST, N

Figure 14. - Isp and thrust Vadeoff at constant pow_.

I

35

24
Report Documentation Page NalJonal Aerorlau_Jcs and

Space Adrfl,rqls[r allorq

1, Report No. 2. Government Accession No. 3. Recipient's Catalog No.

NASA TM-102321

4. Title and Subtitle 5. Report Date

Development of a High Po_'er Microwave Thruster, With a Magnetic

Nozzle, for Space Applications

7. Author(s)

John L. Power and Randall A. Chapman

9. Performing Organization Name and Address

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio 44135-3191

12. Sponsoring Agency Name and Address

National Aeronautics and Space Administration

Washington, D.C. 20546-0001

6 Performing Organization Code

8. Performing Organization Report No.

E-5017

10. Work Unit No

506-42-3 l

11. Contract or Grant No.

13. Type of Report and Period Covered

Technical Memorandum

14. Sponsoring Agency Code

15. Supplementary Notes

Prepared for the 24th Microwave Power Symposium sponsored by the International Microwave Power Institute,

Stantfl)rd, Connecticut. August 21-23, 1989. Randall A. Chapman, NASA/ASEE Faculty Fellow.

16. Abstract

This paper describes the current development of a high-power microwave electrothermal thruster IMET) concep! at the NASA Lewis Research

Center. Such a thruster would be employed in space fi)r applications such as orbit raising, orbit maneuvering, station change, and D_ssibly

trans-lunar or trans-planctary propulsion of spacecraft. The MET concept employs low frequency continuous wave (CW) microwave power to

create and continuously pump energy into a flowing propellant gas at relative high pressure via a plasma discharge. The propellant is heated

to very high bulk temperatures while passing through the plasma discharge region and then is expanded through a throat-nozzle assembly to

produce Ihrusl. as in a conventional rocket engine. Apparatus. which is described, is being assembled at NASA Lewis to lest the MET

concepl to CW power levels of 30 kW at a frequency of 915 MHz, The reich)wave energy is applied in a resonant cavity applicator and is

absorbed by a plasma discharge in the flowing propellant. The ignited plasma acts as a Iossy load. and with optimal tuning, energy absorption

efficiencies over 95 percent (based on the applied microwave power) are expected, Nitrogen. helium, and hydrogen will be tested as

propellants in the MET, at discharge chamber pressures to 10 atm. This facility wilt be the first to test the MFT concept at these power

levels, pressures, and frequency. The performance of the laboratory MET will be enhanced with a strong solenoidal magnetic field in the

thruat-nozzle region, a so-called "magnetic nozzle". The magnetic field acts to compress the ionized propellant along the thruster axis, This

reduces wall impingement and heating and also promotes recombination of the ionic species awa} from the walls, with the consequent

recovery of some of their internal energy. The magnetic nozzle configuration employs a low temperature superconducting magnet with a

maximum field strength of up to 6 T, mounted adjacent to the microwave cavity. This is the first known use of a magne!ic nozzle to improve

the perfol-lnance of an electrothermal thruster operating at high pressure. IApplied magnetic fields have beenlextensively used to enhance the

operation of low pressure magnetoplasmadynamic {MPDI thrusters.) Calculations of the predicted performan_:e of the MET are presented.

"l'hcse calculations are made with a two-dimensional kinetics (TDK) computer code normally used for predicting chemical rocket performance.

]'he code has been extended to include plasma reactions and recently available thermodynamic and kinetic data for hydrogen to temperatures

as high as 20 (R)O K. These are the first known TDK calculations to be made fi_r hydrogen plasmas. The calculatitms indicate that the MET

may reasonabl} achieve a specific impulse of 2(RX) sec tot h_drogen propellant at a pressure of 10 aim, a temperature of 6(X)0 K, and a

power level of 100 kW. The results are in agreement with basic relations for thruster pertbrmance. The specific impulse was found to

increase with increasing chamber temperature and to depend weakly on chamber pressure, while the thrust was found to be directly

proportional to chamber pressure and to depend weakl_ on chamber temperature.

17. Key Words (Suggested by Author(s))

Microwave; Thruster; Electrothermal; Electric

propulsion; Magnetic nozzle; Plasma

19. Security Classif. (of this report)

Unclassified

NASA FORM 1626 OCT 86

18. Distribution Statement

Unclassified- Unlimited

Subject Category 20

T

20, Security Classif. (of this page) / 21. No of pages

Unclassified [ 26

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