Propellant grade hydrazine was prepared with 64 per cent yield and 95.5 per cent purity. Purity of the propellant grade hydrazine was determined using wet chemical, gas chromatographic (GC) and eudiometric methods. It was observed that the compositions containing blends of hydrazine-methyl alcohol-ammonium nitrate and hydrazine-methyl alcohol-ammonium perchlorate were not found to be frozen even after cooling to -65 °C for 30 minutes. Mono and bi-propellant thrusters were designed and developed to demonstrate the performance of prepared propellant grade hydrazine as a promising rocket fuel. Five static tests with 22 N thruster and one static test with 1 N thruster were performed successfully in mono-propellant mode. The hurdles of chamber pressure oscillations were overcome by compact packing of the catalyst. The desired decomposition and chamber pressure were achieved. One static test was performed successfully with 60 N bi-propellant thruster. The desired chamber pressure and thrust were achieved. The combustion was smooth and C* achieved was higher than that of UH-25, N2O4 combination. The performance of prepared propellant grade hydrazine shows it as a promising rocket fuels.


Keywords:   Propellant grade hydrazine,  static test,  mono-propellant,  bi-propellant

Hydrazine is used as fuel for thrusters, in spacecraft propulsion and attitude control systems. Hydrazine thrusters have been widely used in missile and satellite attitude control, velocity correction of spacecraft, and gas generation to provide auxiliary power in aviation because of simplicity and reliability (monopropellant); lowest cost propulsion system (other than cold gas); space storable for long periods (> 12 years demonstrated); low thrust capability and moderate thrust levels available (≤ 400 N)1-4.

In the design of the hydrazine thruster assembly, the collective effects of the injector type, the amount of catalyst packed, and the shape of the catalytic reactor, were the primary concerns5-6. These thrusters have limited life because of voids which ultimately occur due to catalyst loss. The temperature of fuel and catalyst would crucially affect the ignition-delay and induction period of the thruster7-9. Another potential thruster failure mode is poisoning of the catalyst, which reduces the effective area for decomposition. The reduced activity of the catalyst results in a drop in chamber pressure and degradation of the thruster performance due to incomplete decomposition of hydrazine. The decomposition of hydrazine on Shell 405 can be described by the reaction6,10-11. Controlling of the degree of dissociation of NH3 generated from N2H4 decomposition to achieve better characteristic velocities of the thruster, became the basic criterion in the design of injector and catalytic reactor12. Besides temperature, pressure would affect the rate of hydrazine decomposition as well13-14. The hydrazine thruster would acquire better start-up characteristics with small size catalyst, yet a larger pressure drop in the reactor15. In the operation of the hydrazine propulsion system, generally pressure fluctuation was observed. In cold start, a large pressure spike might be generated by excess hydrazine injection and accumulation of hydrazine in the induction period of the decomposition reactions16. Pressure oscillation might also be initiated by valve actuating disturbance and reaction instability.

The performance evaluation of propellant grade hydrazine was tested using self-designed 5 N and 22 N hydrazine propulsion systems at high altitude6,17-18. Low cost catalysts were tested in 1 N, 35 N, and 400 N hydrazine thrusters19. Schmitz describes the qualification test sequence for a long life monopropellant thruster for a satellite reaction control system20. The development of hydrazine monopropellant thrusters carried out at INPE and 400 N monopropellant engine were reported21-23.

However, all these static tests were done with 98.5 per cent pure propellant grade hydrazine. Also with the monopropellant hydrazine thruster, the amount of iridium loading and the percentage of NH3 dissociation aimed are not clearly known. The objective of this work is to demonstrate propellant grade hydrazine in monopropellant and bi-propellant modes with N2O4 as oxidizer and to evaluate thruster performance by studying crucial parameters such as catalyst packing, percentage of NH3 dissociation in monopropellant thruster, and controlling pressure oscillations. In addition, efforts were made to also envisage the feasibility of application of prepared propellant grade hydrazine by formulating few low-freezing hydrazine-based monopropellant blends.

In this work we have prepared propellant grade hydrazine (99.5 per cent) and determined its purity. We have designed and developed both monopropellant and bi-propellant thruster to demonstrate the performance of prepared propellant grade hydrazine as a promising rocket fuel. This paper also discusses about the predictive and static test performances of hydrazine thrusters in both monopropellant and bi-propellant modes.

2.1 Materials

Hydrazine hydrate, 99 per cent LR grade was procured from M/s SD Fine Chemicals Ltd., Mumbai, and Sodium hydroxide, 99.5 per cent AR grade from M/s Finar chemical, Ahmedabad.

2.2 Preparation of Propellant Grade Hydrazine24-26

To 1000 ml round-bottomed flask, 500 ml commercially available 99 per cent pure hydrazine hydrate (64 per cent pure as hydrazine) and various quantities of NaOH (200-415 g) shown in Table 1 were added. The temperature of the water bath was maintained between 65 °C to 85 °C for 1 h to 2 h, a homogenous thick viscous white fluid was obtained. After the completion of the reaction, propellant grade hydrazine was distilled at 65 °C to 70 °C for 2 h under vacuum. The yield obtained varied between 16 per cent -64 per cent, respectively.

2.3 Wet Chemical Analysis

Percentage of hydrazine content was determined by nonaqueous (acetic acid) medium acid base and iodometric (KIO3) titrations1,27-28. The percentage of water content was determined by eudiometric method using the standard procedures29-30

2.4 Gas Chromatography31

All the experiments were carried out in Thermo Quest Trace GC 2000 using packed column using thermal conductivity detector (TCD). Because of the higher polarity of hydrazine, a polar column, 25 per cent PEG on Anakrom (80-100 mesh) was selected and used for all the experiments.

2.5 Determination of Freezing Point at Sub-zero

Tests were conducted for freezing point with sub-zero analysis instrument by taking samples in small glass bottles caped and then sealed with teflon tape. An alcohol thermometer of range -70 °C to +200 °C was placed to check the temperature inside the chamber. The temperature was brought down to -40 °C. The frozen state of the samples were removed and observed. Samples which were not frozen were again placed in the chamber and the temperature was brought down to -65 °C.

2.6 Predictive Performance of Propellant Grade Hydrazine in Mono/Bi-propellant Modes

Predictive performance of propellant grade hydrazine in the monopropellant and bi-propellant modes was determined using hydrazine dissociation reactions and NASA CEA code32, respectively. Heat of formation data calculated from heat of combustion value was used for the theoretical performance evaluation. Performance parameters were obtained with varying air-to-fuel (A/F) ratio.

2.7 Static Test Performance of Propellant Grade Hydrazine

2.7.1 Static Test Performance on Monopropellant Thruster

The stationary rocket combustor was operated at ambient temperature with 1 N and 22 N monopropellant thrusters. The fuel used was propellant grade hydrazine. The pressure measured in the combustion chamber was 7 bar, and at exit nozzle it was ambient. The density of fuel was 1.01 g/cc. The mass flow rate for 1 N was 0.46 g/s and for 22 N thrusters, it varies from 8.62 g/s to 10.5 g/s. The tests were performed with cumulative burn time of 63 s for 1 N thruster, and 23-63 s for 22 N thruster.


2.7.2 Static Test Performance on Bi-propellant Thruster

The stationary rocket combustor was operated at ambient temperature with N2O4 as an oxidizer. The fuel used was Propellant grade hydrazine, at a weight ratio of oxidizer-to-fuel 1.44. The pressure measured in the combustion chamber was 20 bar, and at exit nozzle it was ambient. The density of fuel was 1.01 g/cc. The oxidizer flow rate was 17.7 g/s and fuel flow rate was 12.3 g/s. The test was performed with cumulative burn time of 23 s.

3.1 Preparation of Propellant Grade Hydrazine

Anhydrous hydrazine, used as rocket propellant is obtained by dehydrating hydrazine hydrate. Hydrazine forms an azeotrope with water and ordinary distillation provides hydrazine in the form of hydrate (64 wt per cent of hydrazine). The use of hydrazine in rocket propulsion requires hydrazine to be in the anhydrous form, and hence, the removal of water from hydrate state to produce anhydrous hydrazine is essential to make it suitable for such purposes. The conventional separation techniques for the removal of water experience difficulty as hydrazine forms an azeotrope with water at 71.5 wt per cent of hydrazine. Further, hydrazine and water are highly polar by nature, and there is a strong hydrogen bonding between these. Hence, the combinations of processes are required to seek dehydration24-26.

Table 1. Reaction parameters for preparing monopropellant grade hydrazine

Three experiments were carried with varying ratios of NaOH at different reaction conditions for preparation of propellant grade hydrazine using 99 per cent pure hydrazine hydrate (64 per cent pure as hydrazine). The quantity of NaOH was varied from 200 g to 415 g and time duration was kept between 1 h - 2 h. The reflux temperature was varied between 65 °C to 85 °C. The products were distilled by maintaining the temperature between 60 °C-70 °C. The yields obtained for the three experiments were 16 per cent, 30 per cent, and 64 per cent, respectively. The reaction parameters maintained for these three experiments are given in Table 1.

3.2 Determination of Purity of Hydrazine

The purity of hydrazine was determined using wet chemical method which is based on non-aqueous acid–base titration in acetic acid media and gas chromatographic (GC) method. The total basicity due to hydrazine is calculated using the reaction scheme shown in Eqn (1) and wt per cent of hydrazine will be obtained. From the GC experiments, water content is varied from 10-14 per cent depending on the distillate (S-1, S-2 and S-3) and the method of distillation. However S-4 obtained by distillation of S-3 yielded 99.46 per cent pure hydrazine and 0.54per cent water. No other organic impurities were found in sample. The chromatograms of samples S-3 and S-4 are given in Fig. 1. Water content in the hydrazine determined by the eudiometric method is based on the reaction of water and calcium hydride. Analysis results obtained from wet chemical, GC, and eudiometric methods are given in Table 2.

NH2-NH2+CH3-O||C-OH→NH2-N+H3   O−-C||O-CH3→HClO4CH3-O||C-OH+NH2-N+H3  ClO−4(1)

3.3 Low Freezing Hydrazine-based Monopropellant Blends

A tactical monopropellant should have a wide operating temperature ranging from - 60 °C to + 60 °C. Addition of water to hydrazine lowers the freezing point to forms an azeotropic mixture having boiling point of 120.3 °C and freezing point of - 40 °C. A monopropellant fuel mixture with hydrazine, water and hydrazine nitrate (N2H4 HNO3) had disadvantage of having large percentage of water.

Figure 1. Gas chromatogram of samples S-3 and S-4.

The fuel blend combination chosen for present study are anhydrous hydrazine, methyl alcohol, ammonium nitrate or ammonium perchlorate. Methyl alcohol is miscible with hydrazine and reduces the freezing point of the blend. Methyl alcohol has a carbon content of 37.48 per cent, a hydrogen content of 12.58 per cent and an oxygen content of 49.37 per cent. The substitution of methyl alcohol with water increases the energy value available and makes the blend combination superior. Both ammonium nitrate and ammonium perchlorate have oxidizer as well as fuel part in these; there are the salts obtained from ammonia. This supports the combustion process and produce low molecular weight products during decomposition of hydrazine. Five blends are prepared with varying composition of additives mentioned above by adding propellant grade hydrazine to decrease the freezing point. Tests carried out for analysis of freezing point at sub-zero were shown in the Table 3. It was observed that the composition containing blend 5, i.e., hydrazine and methyl alcohol blend, has frozen. Blend samples 1, 2, and 4 were not frozen at -65 °C even after 30 min whereas blend sample 3 was frozen after 30 min.

3.4 Performance Evaluation of Propellant Grade Hydrazine in Monopropellant Mode

3.4.1 Predictive Performance in Monopropellant Mode

For maximum specific impulse (Isp), it was considered to limit the catalyst length to 26 mm for 40 per cent to 50 per cent ammonia dissociation. All calculations are done for steady-state firing condition of the thruster. Hydrazine purity was considered as 95 per cent, 99 per cent, and 100 per cent, respectively. Various performance parameters such as characteristic velocity (C*), specific impulse (Isp) were determined assuming frozen flow conditions at 7 bar chamber pressure. The parameters were evaluated in chamber, throat and nozzle exit. Supersonic area ratio (Ae/At) was assumed to be 2.75. The characteristic velocity with ammonia dissociation ratio for 95 per cent, 99 per cent and 100 per cent of pure hydrazine was observed as 1284 m/s, 1275 m/s and 1260 m/s, respectively was shown in Fig. 2. The thruster-designed parameters are shown in Table 4.

Table 2. Analysis results of distillate samples

Table 3. Blend formulations for low freezing applications

Figure 2. Characteristic velocity variation with ammonia dissociation.

Table 4. Monopropellant thruster-designed performance parameters for high altitude application


3.4.2 Static Test Evaluation in Monopropellant Mode

The performance of the prepared propellant grade hydrazine (99.5per cent) as rocket propellant in monopropellant mode was investigated by operating stationary rocket motors. Five static tests were conducted with 22 N monopropellant thruster and one test with the 1 N monopropellant thruster in pulse mode at ambient temperature which were designed for sea-level conditions for static testing, and the performance parameters were compared with the designed values.

Table 5 summarized the thruster parameters employed in the tests for monopropellant mode. The monopropellant thrusters Both 1 N and 22 N are made of SS-321 alloy. The shell has a thickness of 2 mm and a factor of safety above 5 for the operating pressure. The injector is a shower head in which the injection tube has a length of 30 mm and is brazed to the injector head. Three injection tubes of 0.58 mm ID are used for 22 N thruster whereas a single tube of 0.19 mm ID is used for 1 N case. The injector and the catalyst chamber are joined via a flange joint using an asbestos gasket for sealing. The pressures are measured upstream of the fuel inlet valve in the fuel tank and at the catalyst chamber exit (entry to the CD nozzle) using Sensotech make strain gauge type transducers.

Table 5. Thruster parameters in monopropellant mode

The motor was operated at a combustion chamber pressure of 7 bar and expanded to ambient pressure of 0.94 bar. Figure 3 represents 22 N thruster with injector. Table 6 represents the test results of granules of iridium-coated catalyst used with alumina. Table 7 provides the details about the amount of catalyst packed for all the six tests. Pulse train was given manually and different pulse widths used for these tests are given in Table 8. Chamber pressure was measured at nozzle entry. The chamber skin temperature reached a maximum of 520 °C with an injector head temperatures less than 160 °C. Chamber pressure oscillations were observed in first test, which was eliminated by proper packing of the catalyst. Table 9 compares the achieved parameters with the designed parameters for all the tests.

Figure 3. Assembly drawing for 22N thrusters.

Table 6. Test results of iridium-coated with alumina granules

Table 7. Catalyst packed in all the tests

Figure 4 represents variation of absolute chamber pressure with time for Test 2. A steady chamber pressure was observed over the 5 s pulse of 7.16 bar. The tank pressure also remained near-constant and the response time of around 90 ms was observed. Figure 5 represents variation of oxidizer manifold pressure with time for Test 6. Steady-state firing was achieved and the tank pressure was steady throughout the test. Response time was around 100 ms. Figure 6 shows variation of nozzle, chamber and injector temperatures with time for Test 6. The maximum skin temperature was observed as 470 °C.

3.5 Performance Evaluation of Propellant Grade Hydrazine in Bi-propellant Mode

3.5.1 Predictive Performance Evaluation in Bi-propellant Mode

Hydrazine purity was considered as 99.5 per cent. Various performance parameters were determined assuming frozen flow conditions at 20 bar chamber pressure. The parameters were evaluated in chamber, throat and nozzle exit. Supersonic area ratio (Ae/At) was taken as 12.5. N2O4 was used an oxidizer for the bi-propellant thrusters. The oxidizer and fuel flow rates are 17.5 g/s and 12.5 g/s, respectively. The thruster-designed parameters are provided in Table 8.

Figure 4. Variation of absolute chamber pressure with time for Test 2 (22N).

Figure 5. Variation of oxidizer manifold pressure with time for Test 6 (1N).

Figure 6. Variation of nozzle, chamber and injector temperatures with time for Test 6 (1N).

Table 8. Designed and achieved parameters for monopropellant test

Table 9. Design parameters of bi-propellant thruster


3.5.2 Static Test Evaluation in Bi-propellant Mode

The performance of the prepared propellant grade hydrazine in bi-propellant mode was investigated by conducting static test using carbon phenolic lined 60N bi-propellant thruster. The ratio of the cross-sectional area of the nozzle exit to-throat cross-sectional area (area ratio) was 12.5. Figure 7 shows schematic diagram of thrusters cross-section used for the test. The bi-propellant thruster is made of HE-15 shell which is internally lined with carbon-phenolic liner for thermal management. The injector is made of Ti-6Al-4V alloy with a triplet injector in O-F-O configuration. The pressures are measured at the upstream locations of both fuel and oxidiser inlet valves and their tanks. The chamber pressure port is located on the injector head. The pressure transducers used are similar to the ones used for mono-propellant thruster testing.

The motor was operated at a combustion chamber pressure of 20 bar and expanded to ambient pressure of 0.94 bar. One static test was performed at ambient temperature. Figure 8 represents variation of absolute chamber pressure, oxidizer manifold pressure, fuel manifold pressure and thrust with time. A single pulse was given for 16 s and a steady chamber pressure of approximately 20 bar and thrust of 56 N was achieved. Combustion instability is not observed. Table 10 represents the performance parameters summary of the static test.

Figure 7. Bipropellant thruster cross-section.

Figure 8. Variation of absolute chamber pressure, oxidizer manifold pressure, fuel manifold pressure, and thrust with time.

(i) Propellant grade hydrazine was prepared in 64 per cent yield with 99.5 per cent purity and determined using wet chemical, GC and eudiometric methods.

(ii) It was observed that the compositions containing blends of hydrazine-methyl alcohol-ammonium nitrate and hydrazine-methyl alcohol-ammonium perchlorate were not found to be frozen even after cooling to -65°C for 30 minutes.

Table 10. Designed and achieved parameters for bi-propellant test

(iii) Five static tests with 22N thruster and one static test with 1N thruster were performed successfully in monopropellant mode. The percentage of dissociation aimed for NH3 is taken as 40 per cent to 50 per cent in the monopropellant thruster. The energetic performance of the thruster matched very closely to the predictive performance with iridium catalyst. Long duration pulse up to 60 s was successfully tested. The hurdles of chamber pressure oscillations were overcome by compact packing of the catalyst. The desired decomposition and chamber pressure were achieved.

(iv) One static test was performed successfully with 60 N bi-propellant thruster. The desired chamber pressure and thrust were achieved. The combustion was smooth and C* achieved was higher than that of UH-25, N2O4 combination.

Both monopropellant and bi-propellant thrusters were designed, fabricated, and validated to demonstrate its capability to meet the demanding mission requirements in reaction control thrusters. The concept of mono/bi-propellant-based thrusters using hydrazine fuel was established. The performance of prepared propellant grade hydrazine shows these as promising rocket fuels.

The authors wish to thank Director, DRDL and Director, Propulsion for supporting this study.

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Mr S. Krishnamachary has completed his MSc Chemistry from Sri Krishna Devaraya University, Anantapur, Andhra Pradesh. Currently pursuing his PhD from DIAT, Pune. Presently he is working as a scientist at Propellant and Polymer Division of Directorate of Propulsion, Defence Research and Development, Hyderabad. His area of expertise is development of liquid propellants for rockets, liquid fuels for ramjets, and thermal protection systems for rocket applications. His contributions includes development of liquid high density hydrocarbon fuels, Mixed oxides of nitrogen (MON-10) for low-freezing applications and silicone-based low density ablative material for external thermal protection.

Dr S. Krishna Mohan is presently working as a scientist at Propellant and Polymer Division of Directorate of Propulsion, Defence Research and Development, Hyderabad. He obtained his PhD in Organic Chemistry from Indian institute of Chemical Technology, Hyderabad. He has received prestigious fellowships from Japan Science and Technology, Japan Society of Promotion of Science and Swiss Postdoctoral Research. He has filed 02 world patents, 06 Indian patents, published over 40 peer reviewed international journal papers and 01 book. His current research interests include development of polymer matrix composites for air frames, ablative materials for thermal protection systems and liquid propellants for rocket and ramjet applications.

Dr S.G Kulkarni is working as Visiting Professor at Defence Institute of Advanced Technology, Pune (India). He has vast experience of teaching at post-graduate level and guided several MTech and PhD theses. He has published a number of research papers in national and international journals of repute and attended seminars / symposia / conferences and presented papers. His areas of interest include high energy materials, detonics, NBC warfares, energetic polymers and nanocomposites. He has delivered a number of invited lectures at various universities, research establishments and ordnance factories in the country.

Dr D. Jayaraman Completed his MSc from School of Chemical Sciences, Pondicherry University. He obtained his PhD from Indian Institute of Technology, Bombay, Mumbai. He joined DRDL (DRDO), Hyderabad at Propellants and Polymer Division under Directorate of Propulsion in 2005. Presently he is working as Scientist-‘D’. His areas of research include synthesis, characterisation and performance evaluation of high density hydrocarbon fuels for SCRAMJET and LFRJ application and materials for thermal protection of rocket motor.

Mr M. Raghavendra Rao is currently working as scientist ‘F’ and is the Head of Liquid Propulsion Division at Defence Research and Development Laboratory, Hyderabad. He has received his BE in Mechanical Engineering from National Institute of Technology, Bhopal, and ME (Aerospace Engineering) from Anna University, Chennai. He is the designer and system manager for ‘Reaction Control Systems’ (RCS) of all the AGNI missile variants. He has more than 20 years of experience in design and developments of various sub-systems for RCS and Velocity Trimming Package (VTP).

Mr L. Dev Singh is presently working as scientist ‘D’ in Liquid Propulsion Division of Defence Research and Development Laboratory, Hyderabad. He has received his BE in Mechanical Engineering and ME in Production Engineering from Osmania University, Hyderabad. He has 10 years of experience in production management in BDL and is presently the In-charge of the liquid rocket engines’ static testing facility at DRDL.

Mr G. Sai Krishna Prasad is currently working as scientist ‘C’ in Liquid Propulsion Division of Defence Research and Development Laboratory, Hyderabad. He has received his BTECH and MTECH degrees in Aerospace Engineering from Indian Institute of Technology Madras. His research areas are design of liquid propulsion systems & development of pulse detonation engine technology.