(c)2000 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

—— — ̂ " -=, _____________ AOO-36899

AIAA2000-3761

Performance Evaluation of an AugmentedHydrazine Thruster

A. OrenRAFAEL, Haifa, Israel

C. GutfmgerTechnion, Haifa, Israel

AIAA/ASME/SAE/ASEE36th Joint Propulsion Conference & Exhibit

16 -19 July 2000Huntsville, Alabama

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(c)2000 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

Performance Evaluation of an AugmentedHydrazine Thruster

Aharon OrenRAFAEL, P.O.Box 2250, Haifa 31021, Israel

Chaim GutfingerFaculty of Mechanical Engineering,

Technion - Israel Institute of Technology, Haifa 32000, Israel

AbstractThe present work describes the development andperformance evaluation of an augmented hydrazinethruster, which provides thrust in the range of 1.0 -0.4N, depending on inlet pressure. The requiredelectrical power is relatively low, i.e. 200 - 400W.An important objective of this work was to find anefficient firing mode that would lead to optimalperformance for given power consumption.

The design goal was to develop and construct athruster with improved performance relative to thatof an ordinary catalytic thruster, and which couldalso be operated during periods of absence of powerwith minimal performance penalty. With this inmind, we have restricted the level of ammoniadissociation to 70 - 80%, thus limiting the perform-ance degradation at zero power to 10% relative tothat of the catalytic thruster), as shown in laboratorytests.

The recommended mode of operation is pulse-train.Each pulse consists of several tens of seconds' offiring, followed by 60 seconds of heating at zeroflow, thus raising the temperature of the heatingsleeve. The above mode yields a higher performancethan that of the common 280 seconds' specificimpulse.

A numerical model was developed in order toanalyze the experimental results and predictexpected performance for further design andoperational modes.

Introduction

Resistojet thrusters combine catalytic hydrazinedecomposition with heating of the decomposedgases by an electrical resistance heater. As a result,the velocity of the gases at the nozzle exit isincreased, leading to enhanced specific impulse. Dueto the higher specific impulse longer missions can be

accomplished or, alternatively, for a specifiedmission the total system weight can be lower.During periods of satellite inactivity the availableelectrical power can be directed to the thruster'sheater. Augmented thrusters (0.3N at O.SkW) havebeen used for North-South station keeping at GEOwith delivered specific impulse of 280s. Animproved model, which delivers 300s, operates inthe Iridium satellite constellation.1

Advances in solar panel and battery developmentduring the last several years resulted in sufficientpower for operating Electrical Propulsion (EP)systems in small satellites too. This allows the use ofEP thrusters for orbit raising and adjustment, fordrag compensation, constellation installment anddeorbit of small satellites that operate at low earthorbits (LEO).The Hall design is the mature one among the plasmabased EP's and is considered as the preferred one2,due to its high specific impulse - 1500s, and long lifeendurance with minimum performance degradation.On the other hand, for power supply under 1 kW thedelivered thrust is in the range of 40 mN.

For a certain value of AV, the required propellantmass can be evaluated by using the Hohmannexpression for orbit transfer. Assuming dry weightof 4 and 18 kg for augmented and Hall thrusters,respectively, the break-even point for mass criteria isreached in the range of 500 m/s.

On the other hand, the augmented thruster has someadvantages, compared with the Hall typeconfiguration, i.e. shorter orbit transfer time andpropellant common with the ACS sub-system.Moreover, the augmented thruster has ratios ofthrust to power far higher than other EP options,because of its high efficiency and modest specificimpulse, its lowest EP system dry mass (becausepower processor is not required) and its clean plume.Those advantages could benefit applications wherepower and thrusting time are limited, (see Sackheim& Byers3)

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(c)2000 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

The present work describes the development andperformance evaluation of an augmented hydrazinethruster, which provides thrust in the range of 1.0 to0.4N, depending on inlet pressure and power supply.The required electrical power is relatively low, i.e.200 — 400 W. An important objective of this workwas to find an efficient firing mode that would leadto optimal performance for given available power.This is achieved by preheating the Heat Exchanger(HX) walls before the gas is fed to the heatingelement.

This study is a continuation of a previous work4 thatfocused on the design of the heating element andparametric evaluation of the thruster designvariables, e.g. ammonia dissociation level andpressure drop. It was found that high rates ofdissociation lead to a high specific impulse due to ahigh amount of hydrogen produced, and lower heatlosses as a result of lower temperatures.

The present design goal was to develop and con-struct a thruster with improved performance relativeto that of an ordinary catalytic thruster, which couldalso be operated during periods of absence of powerwith minimal performance penalty. With this inmind, we have restricted the level of ammonia dis-sociation to 70 - 80% and the pressure drop at theHX. Thus, performance degradation at zero powerhas been limited to 10%, compared to that of thecatalytic thruster, as shown in laboratory tests.

Mission Definition and RequirementsThe thruster should provide the required impulse totransfer a small satellite from an injection orbit to acruising orbit

Design goals:

• Specific impulse higher than 280 s for 400Wheating power.

• Thrust in the range 0.8 - 0.4N, depending oninlet pressure.

• Minimal heat dissipation to the platform -less than 100W.

• The thruster performance for operatingwithout any heating power should be betterthan 90% of that of catalytic thruster.

• The effect of inlet pressure on thrusterperformance should not be excessive.

Design approach

The following design concepts are used in order toimprove performance:

• The overall effectiveness for given power isimproved for reduced mass flow rate. Hence,

a high pressure-drop device should beoutfitted at the thruster entrance.

• Inserting a pressure reducer at the valve inletto reduce the mass flow rate at high pressure.

• The pressure drop across the HX should below in order to maintain high performance forconditions of no heating. This is contrary tothe usual practice of maximizing the heattransfer coefficient by increasing the velocity,which results in higher pressure drop.

• The level of ammonia dissociation should berestricted, since for the case of no power highdissociation moves the working point awayfrom the optimum, thus affecting thrusterperformance.

• It is highly recommended that excess powerbe utilized for preheating the HX sleeves,thus increasing total heat transfer during theflow stage.

ConstructionA cross-section of the thruster is depicted in Fig. 1.The gas generator is based on an existing IN thrusterdesign, but without the catalyst-bed heater. The gasgenerator is connected to the heat exchanger througha short tube of 5 mm in diameter. The selectedflow-control-valve is of the single-coil dual-seattype. This leads to a saving of 100 g in weight ascompared with commonly used double-coil valves.The heating sub-assembly is comprised of aresistance heater, which transfers heat by radiationto the inner surface of the HE. The heat exchanger ismade of two concentric tubes of Niobium C-103. Arectangular helical tunnel is engraved on the innersleeve through which the gas mixture flows.Niobium C-103 is suitable for high-temperatureoperation, and has good electron-beam weldingproperties. Figure 1 also shows the position of thethermocouples used for the experiments describedbelow.

The HE is attached to the heater housing made oftitanium. Titanium has low heat conduction, andallows good welding with Niobium. To reduce heatlosses by conduction, the housing is attached to themounting plate by means of a bracket equipped withan insulation sleeve.

Due to the high temperature of the heating element itshould be protected from oxidization by keeping itin deep vacuum, which is available in space. Duringthe laboratory experiments, the vacuum level in thetest chamber was insufficient, and therefore anevacuated envelope made of titanium was builtaround the heater. This, however, added 500 g to theaugmenting device with considerable absorption ofheat from the hot sleeve, resulting in reduced heatingof the gas.

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The expected mass of the flight model will be in therange of 700 g.

Mode of OperationEvery firing was preceded by a preheating stage toraise the initial sleeve temperature. The results of anensuing long steady state firing test indicated aninitial drop in the temperature of the preheated heatexchanger sleeve, which later stabilized at somevalue. As a consequence, a considerable differenceof AIsp = 20 s existed between the performanceduring the first 10 s and the next 50 s. That point ledus to adopt a pulse-train mode of operation whereduring the time between the pulses the sleevetemperature is raised again, by the heater. In order toobtain the same high Isp for operating at steady orpulse modes one should use 600W for the steadymode instead of 400W required for the pulse mode.

Based on the present study the recommended firingduty cycle, for obtaining Isp higher than 280 s, is20s ON / 50s OFF. The expected total impulse to bedelivered within 1 hour is 600 N-s compared with150 N-s for the Hall thruster.

Discussion of Test ResultsTemperature variations with time during a 200 Wpowered steady firing test, are shown in Fig. 2. Theexternal sleeve reaches a steady state temperature(Ti), which is 200K lower than the initial oneobtained after pre-heating. The gas temperature (T,)exhibits an opposite effect. At the first 20 s it is100K higher than that of steady state, due toexcessive heat transfer because of a highertemperature difference.

The effects shown in Fig. 2 are emphasized in thecase of pulsed firing depicted in Fig. 3. The effect ofpreheating on the temperature of the sleeve (T3) isclearly demonstrated.

The figure indicates that the thermocouple reading atthe nozzle inlet (Ti) does not drop to zero at zero gasflow. This anomaly is due to the thermal inertia ofthe thermocouple material.

Figure 4 shows representative experimental valuesof the specific impulse as a function of inlet pressure(correlated to the mass flow rate) and electric power.

• Curves a and b, compare the Isp of a typicalIN thruster (b) with that of the augmentedmodel (a) operating without additional heating.At low pressure the performance degradation isabout 10%, while at high pressure it is only5%.

• Lines c and e represent the augmentedthruster performance at a high duty cycle, for200 W and 400 W, respectively, with a pulse

train of 60 s firing and 10 s pausing for heating.The maximum Isp reached with 200W was 235s and with 400W it was 270 s.

• Lines d and / represent the augmentedthruster performance at a high heating mode ofoperation (low duty cycle), where the pulsetrain consists of 20 s firing and 60 s pausing forheating. The maximum Isp that was reachedwas 245 s with 200W for case d, and 290 swith 400W for case/

As noted above, the high values of the specificimpulse for cases d and/were achieved because ofreduced accumulated delivered impulse at certainperiods at the expense of operating time. Therequired total time to complete the mission is threetimes longer than required for cases c and e.

Simplified Numerical Model

Sketch No.l: Numerical Model Outline

A simplified numerical model that simulates thelaboratory results, assumes that the gas flow mtakes place in the annular space between thesleeves, while the inner sleeve is subjected to auniform heat flux q", as seen in the sketch above.The following three equations describe the thrusterfiring process:

a) The mass flow rate m through the adiabaticthroat:

m =

b) Pressure drop at the heating chamber :

Pex = Pgg — ———m2

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(c)2000 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

c) Pressure drop at the gas generator:

Pgg = Pin ~ (^tube + ̂ Pcr,n,p + injector + ̂ cat]

The gas temperature at the nozzle inlet is obtainedthrough the transient heat balance model describingthe temperature distribution along the heatexchanger:

d) Heat balance on the inner sleeve:

e) Heat balance on the fluid:

f) Heat balance on the enclosure:

dx-Ten)

2n CTE

The variables are:

th - Mass flow rate (kg/s)Pn - Pressure at nozzle entrance (N/m2)

Pgs - Pressure at gas generator exit (N/m2)

Ts - Temperature distribution on inner sleeve (K)

T - Gas temperature along HX (K)

T - Enclosure axial temperature distribution (K)

Assumptions:

• The catalytic dissociation in the gas generatoris assumed to be 0.8 for all values of mass flowrate. Thermal dissociation is not expected tooccur at the HX, since the gas temperature doesnot reach 1700K.

• The gas temperature at the heater "finger"entrance is assumed to be given by thefollowing expression:

Tln = 940 + 350m - 200 • (l-1.2Pelectr/40Q)(K)

The first and second terms in that equationpresent the measured data for a IN catalyticthruster. The third term stands for cooling of thegas as it flows through the pressure chamber intothe HX. The last term simulates the gastemperature rise due to heat conduction from thehot "finger" to the gas generator body.

The thermal properties of the constructionmaterials and the gas (c , C, k) were taken aslinear with temperature.

Initial temperatures for each operation are300K. The number of pulses is large enough toreach a quasi-steady temperature distribution.

The contribution of the radiation shieldswrapped around the HX "finger" is simulatedby low emissivity on the surface.

The heat transfer coefficients were adjusted toaccount for the area difference between themodel and the real configuration (square spiraltunnel).

The change of nozzle diameter due to thermalexpansion has been taken into account.

The thermal efficiency of the electrical heateris 95%. The remaining 5% power is transferredby conduction through the heater body to theexternal housing.

The specific impulse and thrust are calculatedusing the following expressions :

-* sp ~ ^effective ., | ̂ C p * to

Thrust = mlspg (TV)

(sec)

Numerical AnalysisThe numerical model was calibrated against theexperimental results by inserting effectivenesscoefficients and adjusted heat transfer coefficients.

The model was used for two main purposes:

• To calculate the heat losses at the laboratorythruster, and

• To evaluate expected performance with the flightmodel.

Sketch No.2: Thruster Heat-Balance Scheme

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(c)2000 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

Three main elements of the heat loss are depicted inthe scheme of Sketch No.2 above:

• Conduction to the heater adapter

• Radiation from the hot "finger" to the satelliteenvironment, which is already reduced due tothe assembly of several foils acting likeradiation heat shields, and

• Radiation from the exposed nozzle surface tothe outer space.

The following firing parameters are the basis of thenumerical evaluation for the above heat losses:

- Pre-heating for 300 s.

- Train of 8 20 s ON / 50 s OFF pulses.

- Power is ON during the whole test.

Discussion of Numerical ResultsFigures 5, 6 and 7 summarize the obtained results interms of block diagrams. Each block is built up ofthe above-mentioned three heat loss elements. Thefirst column represents heating with 200W, whilethe second is for 400W. The test was done at inletpressures of 10, 15 and 20 bars. Data for theperformance comparison was taken from the 8th

pulse reflecting quasi-steady values.

Figure 5 shows the calculated values for thelaboratory thruster. The total heat loss (at 10 bar &400W) is 305W where 27% is due to conductionbackwards, 57% is radiation to the satelliteenvironment and 16% is radiation to the outer space.

Calculating the total heat balance on the gas wefound that the net heating power of the gas is 295 W.The "extra" 200W are a benefit of the pre-heatingafter each pulse.

Figure 6 shows the values calculated for animproved laboratory thruster built of a shorter andthinner adapter. The total heat loss (at 10 bar &400W) is 299W, where 21%, 64% and 15% are therelative parts of each contribution. The saving isnegligible, because higher temperatures wereobtained on the external sleeve, which resulted inhigher radiation heat transfer. Therefore, it isessential to improve heat shield performance.

Figure 7 shows the values calculated for aflight-model thruster having an equivalent emissivitythat is by 40% lower than the former one. The totalheat loss was reduced by 20W, resulting in amaximum specific impulse of 300s at 10 bar and400W.

SummaryAn augmented hydrazine thruster was built andtested for steady and pulse mode firing. Theobtained specific impulses were 260 and 280s,respectively. It was shown that sustaining the powerduring no propellant feed, results in a gain of a 20shigher specific impulse. Analyzing the variouselements of heat loss, it was found that radiationfrom the hot "finger" is the dominant one. Withbetter radiation heat shielding the expected specificimpulse can exceed 300s for 400W heating power.

References

1. R. R. Stephenson, "Electric PropulsionDevelopment and application in the UnitedStates", International Electric PropulsionConf, Paper 95-01, Sept. 1995.

2. R. M. Myers, S. R. Oleson, F. M. Curran andS. J. Schneider, "Small Satellite PropulsionOptions", AIAA-94-2997.

3. R. L. Sackheim and D. C. Byers, "Status andIssues Related to In-Space PropulsionSystems", J. Propulsion and Power, Vol. 14,No. 5, September-October 1998.

4. A. Oren, N. Miller and C. Gutfmger,"Augmented hydrazine thruster for smallsatellites" Paper # IAF-94-S. 1.398,Proceedings of the 45th InternationalAstronautical Federation Congress,Jerusalem, Oct. 9-14, 1994.

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(c)2000 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

T1

Fig. 1. General Section of the Laboratory Thruster.

aa>H

800200 250 300 350 400 450 500

Fig. 2 - Temperatures distribution for Steady Firing

550 600Time (s ec)

T! = Temperature of gas at nozzle inlet (Tex)T2 = Temperature of gas generator external sleeveT3 = Temperature of heat exchanger external sleeve

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1500

1400

1300 --

& 1200 - -

| 1100-*-*

« 1000£H 900 -

800 -

700

6000 50 100 150 200 250 300

Time (sec)Fig. 3- Temperature Distribution for Pulsed Train Firing (400W)

18010 11 12 13 14 15 16 17 18 19 20 21 22 23 24

Inlet Pressure (bar abs)Fig. 4-Isp\s. Inlet Pressure at Various Power Levels

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300

250

200

150

100

50

0

15 20Inlet Pressure (bar)

Fig. 5- Laboratory Thruster - Heat Losses

10 15 20Inlet Pressure (bar)

Fig. 6- Improved Laboratory Thruster

*-l space

Eradiation

Econd

400W

20CW

5 10

Fig. 7- Flight Thruster - Heat Losses

15 20Inlet Pressure (bar)