1 k W A m m o n i a Arcjet S y s t e m D e v e l o p m e n t for a Sc ience Miss ion to the M o o n

I E P C - 2 0 0 5 - 0 7 5

Presented at the 29th International Electric Propulsion Conference, Princeton University October 31 November 4, 2005

Dagmar Bock*, Monika Auweter-Kurtz* and Helmut L. Kurtz ++

Universitiit Stuttgart, Institut .fiir Raumfahrtsysteme

Pfaffenwaldring 31, D-~0569 Stuttgart, Germany

bock @irs. uni- stuttgart, de

A b s t r a c t : This p a p e r p resen t s first s teps for t he d e v e l o p m e n t o f a t h e r m a l a rc je t s y s t e m f o r a s c i e n c e m i s s i o n to t he m o o n . A first a p p r o a c h for a p r o p u l s i o n feed s y s t e m as well as resu l t s a c c o m p l i s h e d d u r i n g tes t s w i th a l a b o r a t o r y m o d e l o f t he t h e r m a l a rc je t ARTUR- 2 a r e i n t r o d u c e d . T h e tes t s w e r e d o n e at d i f ferent power levels f r o m 0.9 up to 1 .SkW at mass flows b e t w e e n 10 a nd 15 m g / s h y d r o g e n . T h e c a t h o d e gap was kept c o n s t a n t at 0.8 m m d u r i n g all tes ts . T h e c o n s t r i c t o r d i a m e t e r is 0.6 m m . T h e h ighes t e f f e c t i v e e x c e e d e d

e x h a u s t ve loc i ty was 8 k m / s at a t h r u s t eff iciency of 32%.

N o m e n c l a t u r e

~JT (%) thrus t efficiency Fcot~t (N) cold gas thrus t

~JT,~h (%) theoretical thrus t /~p (s) specific impulse

efficiency rh (rag/s) mass flow

F (N) thrus t Pd (W) electric power

I. I n t r o d u c t i o n

The Ins t i tu te of Space Systems (IRS), University Stut tgar t , launched the "Small Satellite Program" in 2002. In this context four satellite missions are planned: An earth-sensing satellite (Flying L a p t o p l ) , a satellite for technology demonst ra t ion of electric propulsion systems (PERSEUS), a reentry demonst ra t ion satellite (CERMIT 2) and an all electrical satellite mission to the moon (Bw13). In this context all electrical satellite system refers to the propulsion system of the B w l satellite, which will consist of two different parts: a cluster of ins ta t ionary pulsed magne top lasmadynamic thrusters (I-MPDS) 4 and a thermal arcjet thrus ter in the l k W class with ammonia as propellant. Both thrus ter systems are developed and qualified at IRS.

The use of arcjets for Nor th -South s tat ion keeping of geosta t ionary satellites has been state-of- the-art for more than one decade. The new approach is the use of a combinat ion of a cluster of ins ta t ionary pulsed magne top lasmadynamic thrusters SIMP-LEX (Stu t tgar t Ins ta t ionary Magne top lasmadynamic Thrus te r for Lunar Explorat ion) and a thermal arcjet as propulsion unit for a university satellite to the moon.

The flight to the moon is divided into four distinct phases: Phase one starts from G T O and lasts until the spacecraft perigee is raised above the outer van-Allen-belt . Phase two stretches from there to the sphere

*Ph.D. Student t Professor, Head of the Department of "Space Transportation Technology", Institut fiir Raumfahrtsysteme tResearch Engineer, Head of Laboratory @2005 by the authors.

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The 29 ~h International Electric Propulsion Conference, Princeton University, October 31 - November 4, 2005

of the Moon; phase three extends to the stable elliptical lunar capture orbit at an altitude of about 1400 km above the lunar surface and the final phase inserts the satellite into a circular polar mission orbit at an altitude of 100 km around the Moon.

To minimize the t ime satellite spends inside the van-Allen-belt and therefore also to reduce possible radiation damage to the solar panels and electronics on board the satellite, the thermal arcjet thruster, as the "high thrust" propulsion unit (approx. 100 raN), will be used during phase one. As during phase three the forces by the earth, the moon and the sun nearly cancel out the influence of perturbat ions on the satellite orbit increase resulting in a higher demand of thrust during this mission phase. Therefore, the thermal arcjet thruster will be used during this mission phase. During the other mission phases solely the I-MPDS will be used for propulsion.

The thermal arcjet system development includes numerical simulations of the plasma flow, thermal sim- ulations of the thruster, experiments under space environment conditions and the development and space qualification of an appropriate propellant feed system.

II. Propellant Feed System

For the development of an appropriate propellant feed system first studies were accomplished at IRS. The main task of the propellant feed system is to supply gaseous ammonia for injection into the thermal arcjet. Special interest during the first studies was placed on possible difficulties in providing gaseous ammonia inside the complete pipeline system behind the gas generator. This is due to the fact that the propellant will cool down after vaporization and possibly become a mixture of gaseous ammonia and saturated steam. This resulted in the requirement that the ammonia must be gaseous behind the gas generator to guarantee proper functioning of the thruster.

The propellant feed system is designed for a mass flow of 20 - 25 mg/s ammonia. The experience from the development of a 750W thermal arcjet with ammonia as propellant for the satellite mission AMSAT-P3D 5 points at a mass flow between this range. The exact required mass flow for the lunar mission B w l will be defined during the optimization phase of the thermal arcjet thruster. The feeding will be conducted by the pressure difference between the tank and the point of injection into the thruster. The pipelines are made of stainless steel.

As the propellant feed system will be developed and qualified at the Ias , the structure is being kept as simple as possible. A first concept of the propulsion feed system is shown in Fig. 1.

H3Tank

/ J

[ • < • Fill Drain ~ Filter ~ Particle Trap

Check Valve ~ Gas Generator [ ~ Flow Aperture

Control Valve ( ~ Plenum [ ~ Thruster

F i g u r e 1. S k e t c h o f p r o p e l l a n t f eed s y s t e m

A check valve directly behind the tank, where the ammonia is stored as a mixture of gaseous, liquid and saturated steam, opens and closes the propellant flow. The ammonia mixture will be completely vaporized by the gas generator. Two control valves behind the gas generator compensate pressure variations by generating a pressure buffer zone within the volume between them (plenum). Doing so keeps the pressure just in front of the flow aperture within a given range to guarantee the desired mass flow, which is adjusted by a defined

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The 29 ~h International Electric Propulsion Conference, Princeton University, October 31 - November 4, 2005

pressure-ratio over the flow aperture. To guarantee the requirement that only gaseous ammonia is allowed behind the gas generator the tem-

perature and pressure of the ammonia in the pipelines will be measured by pressure transducers and ther- mocouples so that the pipelines can be heated if necessary. Based on this concept the requirements for the different elements of the propellant feed system were defined and a laboratory model is currently being built to perform first system tests.

I I I . E x p e r i m e n t s

The approach to develop a thermal arcjet system for the lunar mission Bw1 is to perform an optimization of an existing laboratory thermal arcjet model with regard to the requirements of the lunar mission. The step after the optimization will be the development of an engineering model under consideration of the optimization results according to the design of the engineering model A a c u s (Fig. 9).

The thruster used for all experiments described here is a modified laboratory model called ARTUR-2 (Arcjet Thruster University Stut tgar t Regeneratively Cooled) with a radiation cooled nozzle. Figure 2 shows the structural design of the thermal arcjet AaTUR-2. The gap between cathode and anode (cathode gap) can easily be adjusted by a screw. The whole thruster can be disassembled so that different nozzle configurations can be tested. To characterize the thruster and its sensitivities the first tests, which are described here, were carried out with hydrogen as propellant. After the characterization and identification of the sensitivities of the thermal arcjet, ammonia will be used as propellant.

The aim of the experiments was to identify the operation parameters of the thruster as a function of mass flow and current. For this reason the mass flow was varied between 10 and 15 mg/s and the current between 10 and 15A. Figure 3 shows the thruster during operation. The cathode gap (0.8ram) and constrictor diameter (0.6 ram) were kept constant during the tests.

Ceramic Insulator Gas Supply Housing 1 Insulator 1 with soldered & Power Connector (rl°l~bdenum) (boron nitrlde)

Cathode Compression Spring Main Insulators Cathode Anode adjustment Onoonol X750) (AI203 Tubes) (Tungsten 2% Th02) (Tungsten 2% Th02)

(stNnless steel)

Figure 2. Structural design of thermal arcjet Figure 3. Thermal arcjet during test

A. E x p e r i m e n t a l S e t u p

At IRS there are two test facilities especially dedicated to developing and qualifying thermal arcjets in the power range between 0.5 and 2 kW. For the development and qualification of the thermal arcjet system for the B w l mission one of the facilities is used for the test of the propulsion feed system and life-cycle tests later in the project.

The test facility as shown in Fig. 4 consists of a cylindrical vacuum chamber made of stainless steel 1.2 m in diameter and 2 m in length. Its own 3-stage roots pumping system with a pumping speed of approx. 13000m3/h provides a background pressure of 10 - 3 hPa prior to the tests and less than 0.2 hPa at thermal arcjet operation with mass flows up to 15 mg/s hydrogen.

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The 29 ~h International Electric Propulsion Conference, Princeton University, October 31 - November 4, 2005

Figure 5 shows the scheme of the thermal arcjet test facility at IRS. The propellant is fed from a bundle of

Ammonia

3 stage Pumping System

Figure 4. Test faci l i ty F igure 5. S c h e m e of tes t fac i l i ty

hydrogen gas bottles or one ammonia gas bott le through pipelines made of stainless steel. The measurement and adjustment of the mass flow is carried out by different mass flow controllers for the different gases (Tylan MFC Typ 280 for hydrogen and Tylan MFC Typ 280S for ammonia). They were calibrated by direct weight measurement with a so-called "sartorius-balance".

The current supply of the thermal arcjet is accomplished by special flight-oriented PCU equipment. It provides a decent current of 0 A up to 25 A and a switch voltage of about 1500 V. The thrust measurement is realized with a pendulum-type thrust stand by measuring the deflection of the pendulum on which the thruster is mounted with a non-contact displacement sensor (NCDT). A set of known weights is used to calibrate the thrust stand previous to every experiment. Optical access to the thruster during the tests is possible through windows inside the tank walls. Control and data acquisition is accomplished by a PC.

B . R e s u l t s

The results presented in this section were achieved during different tests. The graphs and numbers shown here are representative for all the conducted tests. Figure 6 shows the voltage as a function of the current exemplary for the results attained.

The voltage varies between 90 V at a mass flow of l0 mg/s and a current of 12.36 A and 103 V at a mass flow of 15 mg/s and a current of 9.33 A. Here, the characteristic typical for thermal arcjets decreasing voltage with increasing current can clearly be seen. Figure 7 shows a representative result of the thrust vs. input power during different operating conditions.

With the mass flow ~z, the measured thrust F and the electrical power Pd of the thermal artier, the thrust efficiency ~lY can be calculated. Equation 1 gives the equation usually applied for calculating the thrust efficiency. For thermal arcjets of low power, the cold gas thrust has to be considered for the calculation of the thrust efficiency, according to equation 2. 6

F 2 ~l~ - (1)

2#~Pct

F 2 ~l~,~,~ = 2#~Pd + ~ t d (2)

The calculation of the received thrust efficiency for a mass flow of #~ = 12 mg/s and rh = 10 mg/s according to equation 2 and the specific impulse (I~p) according to equation 3 leads to the results presented in Table 1.

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The 29 ~h International Electric Propulsion Conference, Princeton University, October 31 - November 4, 2005

~>

O~

"6 ~>

105.00

103.00

101.00

99.00

97.00

95.00

93.00

91.00

89.00

8.00

i ~ ,~ . . . . ; .... F + 1 0 mg/s - @ ,12 mg/s

- - A - 15 mg/s

I I I

9.00 10.00 11.00 12.00 13.00 14.00 15.00

Current / A

16.00

Figure 6. Voltage vs. current characteristic for hydrogen

Z

2 e- l -

0.16

0.14

0.12

0.10

0.08

0.06

0.04

I I

I I I I

~ ~ - *" '10 mg/s ~ 1 2 mg/s

• 15 mg/s

800 900 1000 1100 1200 1300 1400 1500

P o w e r / W

Thrus t vs. input power characteristic for hydrogen Figure 7.

F I sp -- (3)

#~ go

One can see tha t the characterist ic typical for the rmal arcjets of decreasing th rus t efficiency with increas- ing specific impulse is present. 6

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The 29 th In ternat ional Electric Propuls ion Conference~ Pr ince ton Universi ty, October 31 - November 4~ 2005

Table 1. Measured (F, I, rh) and calculated 01, Isp) performance values

rh (mg/s) I (A) *IT (%) *IT,,~, (%) F (naN) /~p (s)

10 10 34.13 30.26 82.25 816

11 32.44 28.97 82.97 824

12 10 31.16 28.31 89.06 734

11 29.17 26.6 88.04 731

I V . S u m m a r y a n d O u t l o o k

Experiments with hydrogen as propellant at different mass flows between 10 mg/s and 15 mg/s and power ranges between 980 W and 1.6 kW were conducted. The cathode gap was 0.8 mm and the constrictor diameter 0 .6mm. The maximum exit velocity was 8 k m / s at a thrust efficiency of 32 %. The results indicate that the thruster AaTUR-2 shows the typical characteristic of thermal arcjets. Furthermore, a stable ignition behavior was retained during the conducted tests.

For further tests the cathode gap will be varied and different nozzle geometries will be examined. In this context optical tempera ture measurement of the nozzle tempera ture by means of a pyrometer and an infrared camera will be integrated in the tests. As a next step tests will be conducted with anmaonia as propellant to define an optimal working point for the thermal arcjet under consideration of the requirements for the lunar mission B w l .

The development of an engineering model under consideration of the results of the optimization will be following the optimization process of the thermal arcjet AaTUR-2 laboratory model. The design of the engineering model will be based on that of the engineering model of the thermal arcjet AaTus (Fig. 8 and 9).7

I

Figure 8. Structural design of engineering model Figure 9. Picture of engineering model ARTUS

For the flow sinmlation of the thermal arcjet a numerical program called SINA developed at IRS will be used. This program allows the modeling of flows which are of high enthalpy, viscous and turbulent. At present only calculations for hydrogen, argon and nitrogen are possible. 8 For the calculations of anmaonia the chemical processes inside the plasma plume during thermal arcjet operation have to be developed and implemented. The numerical simulations have the tempera ture distribution on the thermal arcjet elements as output data. This will be the input for the thermal modeling. Thermal modeling of the arcjet will be conducted by use of the commercial finite-element-program called ANSYS.

For the thermal modeling first a geometry model inside ANSYS will be built up. The geometry will be modeled as rotat ion-symmetrically constructed for reasons of simplicity.

R e f e r e n c e s

1C-RILLMAYER, Georg; LENGOWSKI, Michael; WALZ, Sebastian; R(SSER, Hans-Peter; HUBER, Felix; von SCH(SNERMARK~ Maria and WEGMANN~ Thomas: Flying Laptop - First Micro Satellite of the University of Stuttgart for Earth Observation and Technology Demonstration, IAC-04-IAA.4.11.P.08, 2004

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The 29 ~h International Electric Propulsion Conference, Princeton University, October 31 - November 4, 2005

2 GRASSLIN, Michael H. and ROSER, Hans-Peter: CERMIT: Controlled Earth Reentry Mini-Vehicle to Improve Technology, IAC-04-V.6.09, 2004

3LAUFER, Rene; AUWETER-KURTZ, Monika; LENGOWSKI, Michael; NAWAZ, Anuscheh; I~()SER, Hans-Peter; von SCHONER- MARK, Maria and WAGNER, Henri: An All Electrical Small Satellite for a Technology Demonstration and Science Mission to the Moon, IAC-04-IAF-Q.2.b.05, 2004

4NAwAZ, Anuscheh; AUWETER-KURTZ, Monika; HERDICH, Georg and KURTZ, Helmut L.: Investigation and Optimization of an Instationary MPD Thruster at IRS, IEPC 05-208, 2005

5AUWETER-KURTZ, Monika; GLOCKER, Bernd; G()LZ, Thomas; KURTZ, Helmut L.; MESSERSCHMID, Ernst W.; I~IEHLE, Martin and ZUBE, Dieter M. : Arcjet Thruster Development, Journal of Propulsion and Power, Vol. 12, No. 6, 1996, pp.1077- 1083

6AUWETER-KURTZ, Monika: Lichtbogentriebwerke fiir Weltraumaufgaben, Stuttgart, Teubner Verlag, 1992 71~IEHLE~ Martin; KURTZ, Helmut L. and AUWETER-KURTZ, Monika : 1 k W Hydrazine Arcjet System Development, IEPC

97-008, 1997 8GRAU, Thomas, Numerische Untersuchung von PlasmawindkanalstrSmungen zur Wiedereintrittssimulatio,~ University

Stuttgart, Diss., 2000

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The 29 ~h International Electric Propulsion Conference~ Princeton University, October 31 - November 4~ 2005