Augmentation of low power hydrazine thrusters

Acta Astronautica Vol. 32, No. 4, pp. 275-281, 1994

Copyright © 1994 Elsevier Science Ltd Printed in Great Britain. All rights reserved 0094-5765/94 $7.00 + 0.00

Pergamon

A U G M E N T A T I O N OF LOW POWER H Y D R A Z I N E THRUSTERS ILAN HADAR and ALON GANY Faculty of Aerospace Engineering and Space Research Institute, Technion--Israel Institute of Technology, Haifa 32000, Israel (Received 25 August 1992; revised 12 May 1993)

Abstract--A general analysis of the augmentation processes for low power hydrazine thrusters, is presented in the light of current concepts of electrical power addition. Among the devices surveyed are the resistojets and the arcjets, which transfer electric energy by physical contact with the hydrazine decomposition gases, and the microwave thrusters which apply electromagnetic fields generated by microwaves. A comparison between the various present and future augmented hydrazine thrusters suggests that the arcjet is the most promising means for the energy addition procedure when both reliability and performance are taken into account.

i. INTRODUCTION'["

The need for improved high performance low-thrust propulsion systems has been driven by the growing performance and mission life time requirements from auxiliary spacecraft engines (e.g. for attitude control and station keeping) in the last 30 years. Figure 1 presents the evolution of the commercial INTELSAT satellite series from 1965 (INTELSAT I - - " E a r l y Bird") until 1987 (INTELSAT VI) [1] in terms of the following characteristics: mission life time, launch fThe Nomenclature is given in the Appendix at the end of this paper. 4

mass, orbital mass, and installed (electric) power. It can be seen that the required masses and power have increased by more than an order of magnitude while mission life time demands have extended from 1.5 to 10 years. Among the propulsion devices installed in satellites for in-orbit maneuvers are chemical thrusters, based on chemical energy producing gasdynamic forces which accelerate the propellant through an exhaust nozzle, and electric thrusters. The latter can be divided into subclasses: electrostatic and electromagnetic engines, based on accelerating a preionized gas by means of electric or electromagnetic fields, respectively; and electrothermal thrusters where ohmic heating is used to elevate the propellent

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ILAN HADAR and ALON GANY

energy, thereafter accelerating the hot gas through a conventional nozzle. One of the most reliable, widespread chemical thrusters, used in spacecraft and satellites for more than 15 years, is the monopropellant hydrazine thruster (MHT). This engine is based on the exothermic decomposition of liquid hydrazine followed by conversion of the thermal energy into kinetic energy of the gaseous products in the exhaust nozzle. The major drawbacks of the MHT is its relatively low energetic performance compared to bipropellant liquid engines. Typical vacuum specific impulse values of the MHT range between 220-240 s, compared to 285-295 s of bipropellant and solid propellant motors [2]. In order to meet the present increasing demands from satellite engines, yet benefit form the well-established, reliable MHT, improvement in the energetic performance of the MHT is essential. Methods of augmenting the MHT performance have recently drawn much attention. The augmented hydrazine thruster (AHT) is a feasible concept combining the advantages of the basic MHT with a higher energetic performance. It is based on the addition of external energy to the decomposition gases of the hydrazine mono-propellant in order to increase the specific impulse. The source of energy for the augmentation process is the available surplus of electric energy, stored in the satellite batteries during non-eclipse operation. As a result, saving of propellant mass and/or increasing mission life time can be obtained. Available power for augmentation is restricted by the spacecraft power system (often solar panels), and is typically about 1 kW for a geosynchronous satellite. Therefore, pronounced augmentation can be achieved particularly in the low thrust (up to 1 N) range. 2. AUGMENTED HYDRAZINE THRUSTER (AHT) CONFIGURATIONS

A schematic block diagram of an AHT is shown in Fig. 2. The liquid hydrazine can undergo thermal or catalytic decomposition, yielding a gaseous mixture containing ammonia, hydrogen and nitrogen at typical temperatures of about 1000 K. The

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Fig. 2. Schematic block diagram of an augmented hydrazine thruster. augmentation process is then introduced, using an electrical source which applies additional heating and elevates the temperature of the mixture. The additional enthalpy is transformed into an increase in the exhaust velocity, thereby increasing the specific impulse. Different electrothermal devices and processes are described below.

2.1. Resistojet The only AHT currently in operation on-board commercial satellites is the resistojet (RJ). It contains the simplest form of heating augmentation: an electric heater. Energy can be added to the gaseous propellant, either internally (by physical contact with a resistor) or externally (by heating the augmentation chamber walls). The first AHT, named High Performance Electrothermal Hydrazine Thruster (HiPEHT), was developed by TRW [3] and serves as a standard engine in the INTELSAT satellites (see Fig. 3). The power augmentation takes place via a vortex heat exchanger. A thrust level of approx. 0.05 N, with a specific impulse value of 305 s have been reported. Another on-board engine (in the SATCOM-7 satellites), which utilizes external heating, is the augmented catalytic thruster (ACT) developed by RRC, U.S.A. [4] (see Fig. 4). Other developments such as the Hughes Electrothermally Augmented Thruster (HEAT)[5] and the Coiled Tube Augmented Thruster (CTAT) developed by ERNO, Germany [6] are in the near-flight stage. RJ performance is discussed later.

2.2. Arcjet The second electrothermal procedure for heat addition to the propellant gases is the arcjet (A J).

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Fig. 4. Illustration of RRC's Augmented Catalytic Thruster (ACT)[4].

The AJ engine is based on an arc discharge, attached between two electrodes (see Fig. 5). The highly ionized arc transfers its electron thermal energy to the gas by means of ohmic heating. The possible use of AJ engines for space propulsion is an old idea [7]. In the early developments (1960s) the use of high power (30-50 kW) AJs was considered for a variety of missions, with no practical success, due to the lack of energy supply sources. Recently, special attention was focused on the development of low power (typically 1-2 kW and 0.2-0.3 Nthrust) engines, particularly for the geosynchronous satellites. Preliminary researches, conducted in the U.S.A. (NASA, LeRC and RRC)[8-11], Japan [12], and Germany [13], were focused on the evaluation of basic characteristics of augmented AJs: start-up capabilities, materials compatibility, thermal loadings, and engine efficiencies. Recently, the extensive work at RRC has led to the qualification

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of a 1.8 kW AJ for a stationkeeping mission on the 1993 Telstar 4 satellite [14,15]. Schematic of RRC's augmented arcjet thruster assembly is shown in Fig. 6. Another low power AJ, currently under development in MBB-ERNO (Germany), is scheduled to be employed on-board the A M S A T P3D staellite in 1995 [16]. 2.3. Microwaves

The direct contact of the heat augmentation device with the decomposition gases, which is typical of the RJ and A J, is associated with inherent problems as regards hardware erosion and overheating. Thus, gas temperature is limited, particularly in the RJ case. Hence, a non-immersed augmentation device would be an attractive possibility. The common use of microwave (MW) ovens, as well as recent developments in plasma physics have led to extensive research concerning M W as part of the electrothermal thruster. An illustration of an A H T based on M W is shown in Fig. 7. Unlike other electrothermal means, where energy is transferred via ohmic resistance, the processes in which energy is transferred to the gas in the M W thrusters are more complicated. With a sufficiently intense M W source the gaseous propellant in the electromagnetic field is ionized and sustained in the plasma state (typical conditions: 1 atm, 3000-4000K). The electrons, accelerated by the electromagnetic field, collide with

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Fig. 6. Illustration of RRC's 1.8 kW augmented arcjet thruster assembly [15].

the heavier molecules, transferring their kinetic energy to form molecules in various excitation levels. A detailed description of the complex chemical processes involved in sustaining the microwave discharge zone can be found in [17]. Downstream, a short recombination section is needed to "thermalize" the energy which is "stored" in the excited species. Consequently, propellants used in the MW thrusters must have the ability to abosrb the MW energy efficiently via a dissociation process, exhibiting, in addition, fast recombination kinetics. Promising working fluid candidates are helium and hydrogen. Recently, a comparison between nonaugmented M W and AJ engines was reported [18], indicating a better overall efficiency for the MW engine with similar working gases (nitrogen and MICROWAVE ENERGY

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2 Fig. 7. Schematic of a microwave electrothermal thruster.

helium), mass flow rates and power levels. Unfortunately, ammonia, the principal decomposition product of hydrazine, has a very low absorption efficiency of MW. This property is related to the unique absorption mechanism of the ammonia molecule in the M W frequency, called "inverse absorption" [19]. This type of absorption exhibits good efficiency only in the liquid phase. Thus, the use of MW heating in the A H T seems not as feasible as the AJ and RJ. Nevertheless, the MW thruster may be more favorable when operating with other propellants. 3. ANALYSIS When operating in a space environment, characterized by no gravity and drag influence in the thrust direction, the increase in the spacecraft velocity (AV) as a result of burning Am mass of propellant is:

) where m 0 is the initial spacecraft mass, g is the earth gravitational acceleration, and Isp is the specific impulse. The impulse (I) added to the spacecraft is: I = ArnI~pg

(2)

Hydrazine thrusters 4. DISCUSSION

Hence, for either A V (stationkeeping) or attitude control maneuver, the most important parameter reflecting the energetic performance of the engine is the specific impulse I~p, which is related to the exhaust gas velocity Uo and the motor thrust F approximately by: Isp~---

U~ F ~ - g rng

For equal thrust levels an increase in the specific impulse implies a proportional decrease in the propellant mass consumption rate [eqn (3)]. Figure 8 illustrates the augmented specific impulse and the fraction of propellant mass which can be saved in an 0.5N hydrazine engine augmented by 1 kW electric power as a function of the power efficiency r/ab and the specific enthalpy absorbed by the hydrazine decomposition gases hab. Since hab---tlabPa,g/in, it is clear that the power efficiency is of major significance for the overall augmentation process. Therefore, the poor absorption efficiencies, which can currently be achieved for MW thrusters, exclude the use of microwaves in the AHT at this time. The AJ and RJ engines, which are based on a similar mechanism of heat transfer to the propellant gas (forced convection and radiation from the hot arc/heater), have similar values of ~/ab" Nevertheless, Fig. 9 which presents the different operational ranges of non-augmented and augmented AJ and RJ thrusters, reveals superiority of the AJ to the RJ reflected by the higher /sp values. The figures consists of theoretical calculations for an H2/N: (2:1 molar ratio) mixture. The limited RJ energetic performance is due to the restriction in allowable temperature of the electric heater ( ~ 2000 K), reflecting current material technology. The AJ can operate at much higher temperatures, thereby yielding higher Isp levels. Typical AJ power efficiency (~/ab) values range from 25 to 35%, as can be seen in Fig. 10, which

(3)

where rh is the propellant mass flow rate. The source for kinetic energy of the exhaust gas is the thermal enthalpy in the combustion chamber, resulting from two contributions: the chemical enthalpy of the propellant without augmentation and the enthalpy gain due to the addition of electric power. Denoting h0 the specific total enthalpy of hydrazine decomposition products without augmentation and P~,g the overall electric power of the augmentation device, the augmented specific impulse is approximately relative to:

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where P~b is the power absorbed by the propellant during the augmentation process.

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Using available low-power energy, the A H T offers a significant reduction in the overall propellant mass of the spacecraft or a parallel lifetime increase. Selection of the augmentation process should take into account both performance and reliability. A technology in which the augmentation process is introduced efficiently by a non-immersed device does not exist currently. Hence, RJ and AJ engines are the only practical A H T choices at this time. The RJ augmented hydrazine thruster has been operated successfully on-board the I N T E L S A T and SATC O M satellites for quite a number of years. However, it offers relatively modest increase in energetic performance (some 30%) over the conventional hydrazine engines. The A J, on the other hand, has very promising augmentation capabilities. Nevertheless, it exhibits much more complex technological problems. Therefore, the recent successful development of a flight qualified augmented AJ engine by R R C opens the door to a new generation of high performance A J - A H T s for various missions. REFERENCES

1. G. Maral and M. Bousquet, Satellite Communication Systems, Chap. 1. Wiley, New York (1987). 2. P. Smith and M. A. Horton, Advanced propulsion systems for geostationary spacecraft--study results. AIAA /SAE/ASME 20th Joint Propulsion Conference, Cincinnati, Ohio, AIAA Paper 84-1230 (1984). 3. S. Zafran, C. K. Murch and R. Grabbi, Flight applications of high performance electrothermal thruster. AIAA /SAE 13th Propulsion Conference, Orlando, Fla, AIAA Paper 77465 (1977). 4. R. T. Fedcona and J. I. Weizman, Satellite reaction control subsystem with augmented catalytic thrusters. AIAA /SAE/ASME 20th Joint Propulsion Conference, Cincinnati, Ohio, AIAA Paper 84-1235 (1984). 5. R. Finston and R. A. Meese, Hydrazine electro-thermally augmented thrust (HEAT). AIAA/SAE/ASME 15th Joint Propulsion Conference, Las Vegas, Nev., AIAA Paper 79-1332 (1979). 6. H. D. Scmitz and M. Steenborg, Augmented electrothermal hydrazine thruster development. J. Spacecraft Rockets 20, 178-181 (1981). 7. R. John and W. Bade, Recent advances in electric arc plasma generation technology. ARS J. 31, 4-17 (1961). 8. S. Knowles, W. Smith, F. Curran and T. Haag, Performance characterization of a low power hydrazine arcjet. 19th AIAA /DGLR/JSASS International Electric Propulsion Conference, Colorado Springs, Colo., AIAA Paper 87-1057 (1987). 9. F. Curran and S. Nakanishi, Low power d.c. arcjet operation with hydrogen/nitrogen propellant mixtures. AIAA/SAE/ASME/ASEE 22nd Joint Propulsion conference, Huntsville, Ala., AIAA Paper 86-1505 (1986).

10. F. Curran and W. Haag, Arcjet component conditions through a multistart test. 19th AIAA/DGLR/JSASS International Electric Propulsion Conference, Colorado Springs, Colo., AIAA Paper 87-1060 (1987). 11. W. Haag and F. Curran, Arejet starting reliability: a multistart test on hydrogen/nitrogen mixtures. 19th

AIAA /DGLR/JSASS International Electric Propulsion Conference, Colorado Springs, Colo., AIAA Paper 87-1061 (1987). 12. T. Yamada, Y. Shimizu, K. Toki and K. Kuriki, Thermal analysis and thrust performance of a low power arcjet thruster. 21st AIAA/DGLR/JSASS International Electric Propulsion Conference. Orlando, Fla, AIAA Paper 90-2581 (1990). 13. M. Glokowski, B. Glocker and H. Kurtz, Experimental investigation of radiation and regenemtively cooled low power arcjet thrusters. 21st AIAA/DGLR/JSASS International Electric Propulsion Conference, Orlando, Fla, AIAA Paper 90-2575 (1990). 14. R. D. Smith, C. R. Roberts, K. Davies and J. Vaz, Development and demonstration of a 1.8 kW hydrazine arcjet thruster. 21st AIAA /DGLR/JSASS International Electric Propulsion Conference, Orlando, Fla, AIAA Paper 90-2547 (1990). 15. W. W. Smith, R. D. Smith, K. Davies and D. Lichtin, Low power hydrazine arcjet system flight qualification. 22nd AIDAA/AIAA/DGLR/JSASS International Electric Propulsion Conference, Viareggio, Italy, IEPC-91-148 (1991). 16. H. L. Kurtz, D. M. Zube, B. Glocker, M. AuweterKurtz, M. Kinnersley, M. Steenborg, G. Matthiius and H. Willenbockel, Low power hydrazine arcjet study. 28th AIAA/SAE/ASME/ASEE Joint Propulsion conference and Exhibit, Nashville, Tenn., AIAA Paper 92-3166 (1992). 17. M. C. Hawley, J. Asmussen, J. W. Filipus, S. Whitehair, C. Hoekstra, T. J. Morin and R. Chapman, Review of research and development on the microwave electrothermal thruster. J. Propulsion Power 5, 703-712 (1989). 18. J. Meuller and M. Micci, Microwave waveguide helium plasmas for electrothermal propulsion. J. Propulsion Power 8, 1017-1022 (1992). 19. C. H. Towense and A. L. Schawlow, Microwave Spectroscopy, Chap. 2. McGraw-Hill, New York (1955). APPENDIX

Nomenclature AHT = AJ = F= g = h= I = Isp = MHT = MW = M~,~ = rn = m0 = Am = P = 1LI = U, = AF = q=

augmented hydrazine thruster arcjet thrust standard Earth gravitational acceleration enthalpy total impulse specific impulse monopropellant hydrazine thruster microwave fraction of propellant mass which can be saved propellant mass flow rate total spacefract mass mass of propellant burned power resistojet exhaust propellant velocity increase in spacecraft velocity efficiency

Subscripts ab = absorbed aug = augmented.