Experimental investigations of the 10 N catalytic hydrazine thruster

Acta Astronautica Vol. 12, No. 4, pp. 237-249, 1985 0094-5765/85 $3.00+ .00 Printed in Great Britain. © 1985 Pergamon Press Ltd.

EXPERIMENTAL INVESTIGATIONS OF THE 10 N CATALYTIC HYDRAZINE THRUSTER

V, SHANKAR and K. ANANTHA RAM Spacecraft Propulsion Systems Division, Auxiliary Propulsion System Unit, Indian Space Research

Organisation, Bangalore-560 017, India

and

K. A. BHASKARAN Thermodynamics and Combustion Engineering Laboratory, Mechanical Engineering Department,

Indian Institute of Technology, Madras-600 036, India

(Received 20 March 1984; in revised form 18 September 1984)

Abstract--Experimental investigations of the 10 N catalytic hydrazine thruster are reported. These thrusters find applications in orbit raising functions of a spacecraft. The hardware was realized and tested in a vacuum chamber (10 -3 mbar vacuum) for its performance. When tested at the design propellant supply pressure of 21.5 bar the thruster developed 10.25 N thrust at an operating chamber pressure of 16.4 bar. The thruster was also tested for off-design conditions (24, 18 and 14.5 bar propellant supply pressures) of operation to determine the steady-state performance. The chamber pressure and vacuum thrust follow more or less a linear law with the propellant-supply pressure. The thruster was also tested for its response characteristics for short (100 ms) firing durations at various propellant-supply pressures (15.5, 18.8, 22.5, 25.6 and 29.5 bar) and the experimental results are reported and discussed. The hydrazine was injected at the room temperature (300 -'s K).

1. INTRODUCTION

Catalytic hydrazine thrusters find extensive applications in attitude and orbit control of a spacecraft. ISRO (Indian Space Research Organisation) Experimental Communi- cation Spacecraft APPLE (Ariane Passenger Payload Ex- periment) and IRS (Indian Remote Sensing Satellite) utilize the hydrazine Auxiliary Propulsion System for their various attitude and orbit control functions. It is also envisaged that a thruster developing about 10 N thrust is required for the spacecraft orbit raising functions (axial thrusting), and it is with the experimental investigations of the 10 N catalytic hydrazine thruster that the present paper is concerned.

Shankar et al. l 1 ] have predicted the concentration of the hydrazine decomposition products along a granular catalytic bed, and have utilized the computer programme developed for the optimal design of the 10 N catalytic hydrazine thruster. The decomposition chamber specifications of 13.0 mm diameter and 16.3 mm length have been arrived at with a bed loading of 35 kg/m2s and a chamber pressure of 15 bar resulting in a 2.5 bar minimum catalyst bed pressure drop and a maximum specific impulse (1852 N s/kg) with respect to the decomposition chamber. The analytical programme has provided us with results concerning the state of the products of the hydrazine decomposition at the end of the decomposition chamber and at entry to the nozzle. Further, based on equilibrium considerations and utilizing a computer programme based on the equilibrium constant approach[2], the state of the products of decomposition as they leave the nozzle is predicted and the details could be found in [3]. Salient theoretical performance results of the 10 N thruster are given in Table 1. The aim and objective of the present paper is to report and discuss the experimental investigations of the 10 N catalytic hydrazine thruster.

2. 10 N THRUSTER ASSEMBLY

A photographic view of the thruster assembly is shown in Fig. 1. The assembly consists of a decomposition chamber wherein the catalyst particles (Shell 405 gran- ules - 2 0 + 30) are packed. The propellant supply is through a single element capillary injector.The inner and outer diameter of the capillary are 0.6 mm and 1.2 mm, respectively. A conical convergent-divergent nozzle is used for expansion of gases to produce the required thrust. A mounting flange facilitates the mounting of the thruster to the test stand or thrust balance, and an interface flange is used for mounting the thruster to the ON/ OFF valve through a " O " ring seal. A thermal standoff keeps the ON/OFF valve at a safe distance from the decomposition chamber, and pickup adaptors are used for the temperature and the pressure measurements. The pickup adaptors are welded to the decomposition chamber through the TIG (Tungsten Inert Gas Welding) pro- cess whereas other parts are vacuum brazed using gold- nickel wire.

A cartridge heater could be mounted on the heater- mount to maintain the catalyst bed start temperature at 363 K during short duration firings. The thruster parts are made of high temperature nitride resistant alloy. The thruster assembly weighs 82.3 g. The thruster is provided with ON/OFF valve for test purposes. The ON/OFF valve developed in ISRO is a coaxial type of valve, that is a valve in which the solenoid is wound round the flow path of the valve and the design is soft seated. The valve is capable of 2.5 million cycles of wet operation while maintaining the performance requirements.

3. EXPERIMENTAL SETUP

The experiments were carried out in a vacuum chamber equipped with appropriate systems for feeding the

238

Table 1. Theoretical performance of the 10 N thruster[3]

State of the Products of Hydrazine Decomposition at the End of the Decomposition Chamber:

Temperature (K) Mole Fractions:

Ammonia Nitrogen Hydrogen Hydrazine

State of the Products of Decompo- sition at the Exit of the Nozzle:

Temperature (K) Mole Fractions:

Ammonia Nitrogen Hydrogen Hydrazine

Nozzle Exit Plane Pressure (mbar) Characteristic Velocity (m/s) Vacuum Specific Impulse (N s/ kg) Equilibrium Expansion Frozen Expansion Thrust at Design Conditions of Operation (N) Design Propellant Supply Pressure (bar) Decomposition Chamber Pressure Decomposition Chamber Diameter (mm) Decomposition Chamber Length (mm) Nozzle Throat Diameter (mm) Nozzle Area Ratio

1245

0.48 0.24 0.25 0.03

398

0.95E-07 0.4 0.6 0.24E-21

30 1268

2050 2032

10

21.5

15.0 13

16.3

2.25 50

V. SHANKAR e t al.

A schematic of the experimental set-up is shown in Fig. 2. The regulated supply of the pressurant, namely, gaseous nitrogen, was obtained from the nitrogen cyl- inders. The gaseous nitrogen conformed to the specifications given in [5]. In Fig. 2 are also indicated locations of the various transducers and the salient measurements performed included:

(a) Pressures using 21 NS ISRO make strain gauge type transducers.

(b) Propellant flow rate measured using a positive displacement type flow meter supplied by M/s ERNO of West Germany[6].

(c) Thrust measured using a thrust balance incorpo- rating soft mounting flexures and LVDT (Linear variable differential transformer).

(d) Temperatures sensed by Chromel-Alumel ther- mocouples.

(e) The exhaust gas analysis was carried out with Masslyser[7].

The details of measurements are given in Table 2. All data were recorded on a UV (ultraviolet) recorder.

4. 10 N THRUSTER HOT FIRING

The thruster under present investigations was sub- jected to hot firing in a vacuum chamber (1200 mm diameter x 1200 mm long) capable of 10 3 mbar initial vacuum. To determine the response characteristics and the steady-state performance characteristics of the thruster the following tests were carried out.

propellant hydrazine and recording the output of various transducers. The monopropellant-grade hydrazine man- ufactured in ISRO was used for testing the thruster. The hydrazine conformed to the specifications' given in [4]. The vacuum system has the capability of 10 _3 mbar initial vacuum, and a vacuum of better than 5 mbar could be maintained during steady-state firing of the thruster.

4.1 Short duration ( I O0 ms) firing The ON/OFF valve was tested under various propel-

lant supply pressure conditions for its response characteristics. The opening and closing response values were better than 10 ms.

The cartridge heater (2 W) was switched on as soon as the initial vacuum level (10 -3 mbar) was built up. The steady temperature of the decomposition chamber

Fig. 1. 10 N thruster assembly. 1: Nozzle. 2: Decomposition chamber. 3: Pickup adaptors. 4: Cartridge heater mount. 5: Thermal standoff. 6: Mounting flange. 7: ON/OFF valve interface flange.

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Table 2. Details of measurements

SI. Type of Range of Accu- No. Parameter Instrument Measurement racy Remarks

1 Propellant supply pres- 21 NS ISRO 0-30 bar ---1% sure, P, strain gauge

transducer 2 Upstream decomposition 21 NS ISRO 0-30 bar -+ 1% Diaphragm water cooled

chamber pressure, P2 strain gauge transducer

3 Downstream decomposi- 21 NS ISRO 0-30 bar ± 1% Diaphragm water cooled tion chamber pressure, strain gauge P3 transducer

4 Propellant flow rate, th Positive dis- 500-6000 - + 2 % Limitation--Pressure 24 bar placement type mg/s flow meter

5 Thrust, F Thrust balance 0-12 N ---2% 6 Decomposition chamber Chromel-Alumel 278-900 K ±2%

body temperature, T~ thermocouple 7 Temperature of the gas- Chromel-Alumel 298-1300 K -+2% Glass to metal seal in the

eous products of the hy- thermocouple ferrule of the adaptor to pre- drazine decomposition at vent gas leakage upstream decomposition chamber location, /'2

8 Temperature of the gas- Chromel-Alumel 298-1300 K - 2% eous products of the hy- thermocouple drazine decomposition at downstream decomposition chamber location, T3

9 Gas temperature at the Chromel-Alu- 298-500 K -+2% nozzle exit plane, T4 mel

thermocouple 10 The exhaust gas Masslyser -+ 1%

composition 11 Vacuum level Vacuum gauges 10-3-5 mbar +-5%

Glass to metal seal in the ferrule of the adaptor to pre- vent gas leakage

body recorded was 363 K. The thruster was fired for a pulse duration of 100 ms at five different chosen propellant supply pressures, namely, 15.5, 18.8, 22.5, 25.6 and 29.5 bar. The vital measurements for these tests were the propellant supply pressure (P~), decomposition chamber body temperature (T~), which is indicative of the catalyst bed conditions, and the downstream decomposition chamber pressure (P3). The recordings were obtained at a high chart speed to facilitate easy reading of the response values.

4.2 Steady firing to measure the gas temperatures To measure the temperatures of the gaseous products

of the hydrazine decomposition the thruster was fired for 50 s duration. The vital measurements in this test were the upstream decomposition chamber temperature (/'2), the downsteam decomposition chamber temperature (7"3) and the exit gas temperature (/'4). The propellant supply pressure (P1) was set to 21.5 bar corresponding to the design conditions of operation.

4.3 Steady firing to determine the performance at the design conditions of operation

To verify the steady-state performance characteristics of the thruster, it was fired for a duration of about 2 s at the design propellant supply pressure of 21.5 bar. This test was carried out after ensuring that the decomposition chamber body temperature (T,) indicated a steady value (about 950 K), thereby ensuring steady operating con-

ditions. The vital measurements were the propellant supply pressure (P,), upstream decomposition chamber pressure (P2), downstream decomposition chamber pressure (P~), thrust (F) and propellant flow rate (rh).

The nozzle exit plane pressure was 30 mbar for the design conditions. A vacuum of better than 5 mbar was maintained in the chamber during steady firing. Hence, the nozzle could be expected to flow full, thereby meet- ing the primary test requirement to determine the steady- state performance characteristics.

4.4 Steady firing to determine the performance at the off-design conditions of operation

The flow measurement setup utilized for evaluating the steady-state performance of the thruster had a limitation that the propellant supply pressure cannot exceed 24.0 bar. Hence, the thruster was tested for 24.0 bar (above the design propellant supply pressure of 21.5 bar) and for 18.5 bar and 14.0 bar (below the propellant supply pressure of 21.5 bar). Beyond 24.0bar propellant supply pressure no flow measurements could be carried out. Hence, the range of testing was 14.5-24.0 bar propellant supply pressure to experimentally determine the performance of the thruster during the off-design conditions of operation. The vital measurements were once again the propellant supply pressure (P~), upstream decomposition chamber pressure (P2), downstream decomposition chamber pressure (P3), flow rate (rh) and thrust (F).

Catalytic hydrazine thruster 241

Fig. 3. 10 N thruster mounted on thrust balance inside vacuum chamber. 1: Vacuum chamber. 2:10 N thruster assembly. 3: ON/OFF valve. 4: Thrust balance. 5: Water cooled pressure pickups. 6: Propellant supply pressure

pickup. 7: Chromel-Alumel thermocouple.

4.5 Steady firing for the gas sample collection In the present investigations a low pressure gas col-

lection system was employed. This essentially consisted of the sample probe located at the nozzle exit plane of the thruster mounted inside the vacuum chamber, a man- ually operated valve and a gas sampling bulb with inlet and outlet stopcocks situated outside the vacuum chamber. The exhaust gas sample was collected during steady firing (about 2 s) and analysed in the Masslyser. The test was carried out at the design propellant supply pressure of 21.5 bar and it was ensured that the gas sampling bulb was also evacuated before collecting the sample.

4.6 Photographic view of the test setup A photographic view of the 10 N thruster assembly

mounted on the thrust balance and inside the vacuum chamber is shown in Fig. 3.

5. RESULTS AND DISCUSSION

The 10 N thruster hardware was tested for short duration firings of 100 ms to evaluate the response characteristics and steady-state firings of about 2 s, which were carried out for evaluating the steady-state performance characteristics. The experimental results are presented and discussed below.

5.1 Response characteristics of the thruster For investigating the response characteristics of the

thruster the buildup and the decay of the downstream decomposition chamber pressure (P3) are important. The thruster was fired for a duration of 100 ms (one pulse) at five different propellant supply pressures (P0, namely, 15.5, 18.8, 22.5, 25.6 and 29.5 bar. The decomposition chamber body temperature (T~) during the start of each pulse was about 363 K. Results of the tests are discussed below.

Figure 4 gives a typical pressure trace obtained during 100 ms vacuum firing at the propellant supply pressure of 18.8 bar. The transients that describe the response characteristics are indicated in the figure. The ON/OFF valve (solenoid actuated) current trace is also shown in the figure.

5.1.1 ON~OFF valve opening response. This is the time interval which elapses between the electrical pulse supply signal and the position of attainment of the fully open valve condition. The ON/OFF valve opening response was found to be better than 10 ms.

5.1.2 ON~OFF valve closing response. This is the time interval which elapses between the electrical supply closing signal and the position of attainment of the fully closed valve condition. As observed in the figure there is a steep fall in the current trace on the valve closure signal. Hence, the valve closing responses could not be measured using this trace. However, through appropriate electronic circuitry the valve closing response was found to have a value better than 10 ms.

5.1.3 Thruster' 'ON" response. This is the time elapsed from the valve signal open to 90% average P3.

5.1.4 Thruster "OFF" response. This is the time elapsed from the valve signal close to 10% average P3.

5.1.5 Ignition delay. This is the time taken from the valve signal open to buildup 1% average P3.

In Fig. 5 are presented the variations of thruster ON response and OFF response with variations in the propellant supply pressure (P~). It could be observed from the figure that when the propellant supply pressure (P0 increases from 15.5 to 29.5 bar, the OFF response of the thruster decreases from 102 to 55 ms. It is observed that the OFF response is always higher than the ON response. This is because there is a certain amount of propellant getting entrapped in the capillary tube between the ON/OFF valve outlet and the decomposition chamber inlet (known as holdup volume). The holdup volume is

242

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V. SHANKAR et al.

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kept at a minimum and it is 20 mm 3 in the 10 N thruster capillary injector. The holdup volume propellant entering the decomposition chamber slows down the pressure decay. The effect of increasing the propellant supply pressure is to decrease the response values. The observed ignition delay was about 17 ms and remained more or less constant with the propellant supply pressure.

5.2 Thruster steady-state performance characteristics Performance characteristics of the 10 N thruster are

evaluated during steady-state firing (about 2 s) at four different propellant supply pressures (P,), namely, 14.5, 18.0, 21.5 and 24.0 bar. Results of the tests are tabulated in Table 3 and these are discussed below.

5.2.1 Decomposition chamber pressures. Figure 6 gives

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Table 3. 10 N thruster steady-state firing data supply pressure (P0. For the design propellant supply

Parameters Pressure drop in the decomposition chamber, (P2-P3) (bar) Decomposition chamber pressure roughness (%) Vacuum thrust, F (N) Propellant flow rate, rh (mg/s) Vacuum specific impulse, l,p (N s/kg)

Propellant Supply Pressure, P, (bar) 14.5 18.0 21.5 24.0 2.8 2.5 2.3 2.1

12.46 10.0 8.6 7.4

5.5 7.9 10.25 11.9 3100 3950 4850 5450

1774 2000 2113 2183

the downstream decomposition chamber pressure (P3) traces obtained during steady-state firing (about 2 s) in the vacuum chamber, whereas, Fig. 7 gives the upstream decomposition chamber pressure (P2) traces. Both downstream (P3) and upstream (P2) chamber pressures measured are plotted in Fig. 8 to find out how they vary with the propellant supply pressure (Pt). Both upstream (P2) and downstream (P3) decomposition chamber pressures measured increase with an increase in the propellant

pressure (P~) of 21.5 bar the downstream decomposition chamber pressure (P3) is 16.4 bar, whereas the measured upstream decomposition chamber pressure (P2) is 18.7 bar and, hence, there is a pressure drop of 2.3 bar in the packed bed. When the propellant supply pressure (P:) varies from the design value both upstream (P:) and downstream (P3) decomposition chamber pressures follow more or less a linear law when plotted against propellant supply pressure (P0. The maximum and minimum downstream decomposition chamber pressures (P3) recorded were 18.2 bar and 11 bar, respectively, against the propellant supply pressure (Pj) of 24 bar and 14.5 bar, whereas the maximum and minimum upstream decomposition chamber pressures (P2) recorded were 20.3 bar and 13.8 bar.

5.2.2 Chamber pressure roughness. Schmitz[8] has defined the chamber pressure roughness as the average maximum pressure minus the average minimum pressure divided by the mean pressure (P3). When tested for the design propellant supply pressure of 21.5 bar it was observed that the roughness experienced by the 10 N thruster was 8.6%.

The chamber pressure oscillations are also dependent

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on the ratio of pressure drop in the injector to the chamber pressure. It is a general practice in rocketry to allow for an injector pressure drop of 15-20% of the chamber pressure. When the thruster is tested for various propellant supply pressures (24-14.5 bar) it is operated in a blow-down mode. It is an inherent characteristic of this type of operation that the injector pressure drop as a percentage of chamber pressure varies with the propellant supply pressure. An injector pressure drop of 20.3% of steady state chamber pressure was obtained when the 10 N thruster was tested at 24 bar propellant supply pressure, whereas an injector pressure drop of 6.36% of steady-state chamber pressure only was obtained at 14.5 bar propellant supply pressure. This has resulted in a decrease in roughness from 12.46% to 7.4% when the propellant supply pressure (P~) was increased from 14.5 to 24 bar and the variation of roughness with the propellant supply pressure is plotted in Fig. 9.

5.2.3 Steady-state thrust. Thrust values measured during vacuum firing are indicated in Table 3 and a plot of the variation of thrust with the propellant supply pressure is shown in Fig. 10. Thrust (F) increases with the increase in the propellant supply pressure (P0.

At the design propellant supply pressure (Pt) of 21.5 bar vacuum thrust developed by the 10 N thruster was 10.25 N. The maximum and minimum thrust developed by the thruster were 11.9 and 5.5 N against the propellant supply pressures of 24 and 14.5 bar, respectively. Vac- uum thrust more or less follows a linear law with the propellant supply pressure~

5.2.4 Temperature of the gaseous products of hydrazine decomposition in the decomposition chamber and at the nozzle exit plane. These were experimentally measured during steady-state firing (about 50 s) for the design propellant supply pressure of 21.5 bar. The maximum steady temperatures recorded were 1207, 1209 K at upstream (T2) and downstream (T~) decomposition chamber locations, respectively. A difference of only about 2 K was observed between upstream and downstream temperature sensors. This is rather small as the sensors are only 15 nun apart and the equilibrium temperature is attained within a very small axial distance of the decomposition chamber. A temperature of 376 K was recorded by the temperature probe (/'4) at the nozzle exit plane, thereby indicating that the gaseous products of hydrazine decomposition are exhausted from the nozzle at the temperature of 376 K.

5.2.5 Exhaust gas composition analysis. The exhaust gas samples were collected during steady-state firing (about 2 s) in the vacuum chamber. A low-pressure gas collection system was employed for collecting the samples, and the propellant supply pressure (P~) was set to 21.5 bar corresponding to the design conditions of operation. The samples were then analysed using a mass spectrometer (Masslyser). The sample spectra is shown in Fig. 11. The mass numbers are also indicated in the spectra. The major constituents of the sample were hydrazine, nitrogen and ammonia, and in terms of mole fractions: Ammonia = 0.057, Nitrogen = 0.4, Hydro- gen = 0.514 and Hydrazine = 0.029.

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PRESSURE

Fig. 11. Sample spectra.


5.3.4 Performance calculation based on the percentage ammonia decomposition (X). The performance of a catalytic hydrazine thruster may be determined by examining the characteristic velocity (C*). However, there is another single parameter unique to the catalytic

N hydrazine thruster to completely specify thruster performance. This parameter pertains to the amount of ammonia remaining in the decomposition products as the gases leave the catalytic bed and is termed percentage A

IL ammonia decomposition (X). The reactions are assumed to cease after the gases have left the catalytic bed, thus freezing the chemical composition of the gases and fixing the performance of the thruster[9, 10]. The mean molecular weight of the combustion products can be calculated as:

5.3 Performance derived from measurements 5.3.1 Specific impulse. The specific impulse is an

important performance parameter that can be derived once the propellant flow rate and thrust are known:t

F lsp = _ . (1)

m

The experimentally determined propellant flow rate (rh) was 4850 mg/s for the design propellant supply pressure (P~) of 21.5 bar. Hence, the vacuum specific impulse (1,~) is 2113 N s/kg. Vacuum specific impulse data for various propellant supply pressures are given in Table 3.

The variation of vacuum specific impulse (lsp) vs. propellant supply pressure (P~) is plotted in Fig. 12. The maximum and minimum specific impulse developed by the thruster are 2183 and 1774 N s/kg against the propellant supply pressure (P~) of 24 and 14.5 bar, respectively, and the specific impulse (Isp) increases with an increase in the propellant supply pressure (P~).

5.3.2 Characteristic velocity (C*). Another performance parameter that can be derived from measurements is the characteristic velocity (C*) and is given by:

- - 9 6 M - - - - - - 7 . ( 5 )

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It is observed from the measured species composition that the molecular weight of the gaseous products of the hydrazine decomposition works out to 14.1 kg/kg mole, and hence the percentage ammonia decomposition (X) works out to 45%.

Temperature at the end of the catalytic bed is calculated as:

T 3 --- 1650 - 750 X. (6)

This works out to 1312 K. This is higher than the measured temperature of 1209.16 K at the downstream decomposition chamber location. This is due to the fact that eqn (6) gives the adiabatic reaction temperature. The specific heat ratio (?) can be calculated as:

? = 1.15 + 0.22X. (7)

This works out to 1.249. The vacuum specific impulse (lsr) and the character-

istic velocity (C*) can be worked out using the following relations:

The C* value for the 10 N thruster works out to 1319 m/s at the design conditions of operation (propellant supply pressure = 21.5 bar).

5.3.3 Vacuum thrust coefficient (CI). The value of the thrust coefficient is given by the following equations:

lsp

F Cf = AtP3.

The value of the thrust coefficient based on experimental measurement works out to 1.6 for the 10 N thruster and at the design conditions of operation (propellant supply pressure = 21.5 bar).

tNomenclature is given in Appendix at end of paper.

I~ e = ( {1 - (P,/P3) (:-''.'} , (8)

k M7 _1 (9)

(3) The specific impulse works out to 2188 N s/kg thus more or less matching with the value of 2113 N s/kg based on thrust and the propellant flow rate measure-

(4) ments. The characteristic velocity is 1329 m/s against 1319 m/s based on measurements of chamber pressure and propellant flow rate.

6. C O N C L U S I O N

The 10 N catalytic hydrazine thruster has been studied experimentally. This thruster has the potential of finding practical application in the orbit raising function (axial

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DATA,,

thrusting) of a spacecraft. The thruster could also find application in attitude and orbit control applications of, say, 1500 kg class geosynchronous spacecraft having a life-time of about seven years. Inclination corrections, meaning station keeping operations, could also be carried out using these thrusters by suitable orientation.

The design propellant supply pressure is 21.5 bar. The thruster could be used in a blow-down mode when the propellant supply pressure varies from 24 bar at BOL (beginning of life) to 14.5 bar at EOL (end of life) of a spacecraft. The response characteristics and the steady- state performance characteristics have been obtained experimentally.

Further, it is planned to test the thruster for various duty cycles (5 to 98%) with ON time varying from 10 ms to 500 ms and for about two lakh pulses. Also steady firing of about 5000 s is planned. It is also planned to demonstrate the cold start (bed temperature 5 C) capa- bilities of the thruster when studies on the start characteristics at different conditions could be made. The thruster is expected to undergo about 30 kg hydrazine throughput.

Acknowledgements--The authors are thankful to Dr. A. E. Mu- thunayagam, Programme Director, Auxiliary Propulsion Sys- tem Unit, Indian Space Research Organisation, for encouraging studies of catalytic hydrazine thruster. The authors wish to thank the authorities of Indian Institute of Technology, Madras, for encouraging publication of the work relating to the first author's Ph.D. registration under the external registration scheme.

REFERENCES

1. V. Shankar, K. Anantha Ram and K. A. Bhaskaran, Pre- diction of the concentration of hydrazine decomposition

products along a granular catalytic bed. Acta Astronautica 11, 287-299 (1984).

2. T. K. Bose, High temperature gas dynamics, chapter VI- Equilibrium Composition of a Reacting Gaseous Mixture, pp. 136-137. Macmillan India Ltd. (1979).

3. V. Shankar, Analytical and Experimental Studies of 10 N Catalytic Hydrazine Thruster, Ph.D. Thesis. Thermody- namics and Combustion Engineering Laboratory, Mechan- ical Engineering Department, Indian Institute of Technol- ogy, Madras (1983).

4. Military Specification Propellant Hydrazine, MIL-P-26536C, Amendment I, July (1974).

5. Military Specification, Propellant Pressurising Agent, Ni- trogen, MIL-27401.

6. Flow Measurement System for Hydrazine Engines, ERNO, Raumfahrttechnik, GmbH, PDM S N2 (1983).

7. Masslyser MS001 Model, Hind Hivac, Bangalore, India (1983).

8. Schmitz, Long life monopropellant hydrazine engine development programme. AFRPL TR-71-103 (1971).

9. B. W. Schmitz and W. W. Smith, Development of Design and Scaling Criterion ]br Monopropellant Hydrazine Re- actor Employing Shell 405 Catalyst, Report 66-R-76 Vol. II. Rocket Research Corporation, Seattle (1966).

10. D. Adler, E. Dubrov and Y. Manheimer-Timnat, The performance of a hydrazine engine with an improved catalyst, Acta Astronautica 2~ 613-625 (1965).

APPENDIX

Nomenclature A area of cross section, m 2

C* characteristic velocity, m/s C/ vacuum thrust coefficient F thrust, N g acceleration due to gravity, m/s 2

1~ vacuum specific impulse, N s/kg M mean molecular weight of the hydrazine decomposition

products, kg/kg reel rn propellant mass flow rate, mg/s P pressure, bar

Catalytic hydrazine thruster

P~ propellant supply pressure, bar P2 upstream decomposition chamber pressure, bar P3 downstream decomposition chamber pressure, bar Pe exit pressure, mbar R universal gas constant 848.39 kg m/kmol K T temperature, K

Tt decomposition chamber body temperature, K T2 temperature of the gaseous products of the hydrazine

decomposition at the upstream decomposition chamber location, K

T~ temperature of the gaseous products of the hydrazine

249

decomposition at the downstream decomposition chamber location, K

7"4 temperature of the gaseous products of hydrazine decomposition at the nozzle exit plane, K

X percentage ammonia decomposition, %

Greek y ratio of specific heat of gases

Subscript e exit condition t throat condition

AA 12:4-D

Documents

Experimental investigations of the 10 N catalytic hydrazine thruster