With the increasingly urgent demand for non-toxic, high-specific-impulse propellants in space propulsion systems, ammonium dinitramide (ADN)-based liquid propellants have attracted extensive attention as a new-generation green alternative to hydrazine-based fuels. However, existing ADN-based thrusters all adopt the catalytic ignition technology route, which suffers from inherent drawbacks: the catalyst cannot withstand temperatures exceeding 1500 K, long preheating times are required prior to startup (in the Prisma mission, in-flight preheating lasted 600–720 s with a single-event energy consumption of 25 kJ), and insufficient preheating may lead to "hard start" or even explosion—these issues severely constrain the rapid response capability and operational safety of the thrusters. Although active ignition methods such as resistive ignition and laser ignition have been validated at the single-droplet level, how to achieve catalyst-free electric ignition at the thruster system level and systematically evaluate its combustion characteristics remains a critical engineering problem urgently to be addressed in the field of green space propulsion.
In a recent study published in Space: Science & Technology , a research team from Beijing Jiaotong University proposed an ADN-based thruster employing a combined scheme of resistive ignition and arc-assisted combustion, verifying for the first time the feasibility of electric ignition technology at the thruster system level. The study innovatively designed a multi-layer honeycomb decomposition electrode structure to enlarge the contact area between the propellant and the electrodes, and conducted systematic hot-fire tests under the conditions of 5 N thrust and a propellant mass flow rate of 2.5 g/s. The effects of ignition voltage, arc loading time, electrode gap, and electrode orifice diameter on thruster performance were investigated. The results show that under the operating conditions of 80 V ignition voltage, 3 mm electrode gap, and 0.8 mm electrode orifice diameter, the thruster achieved cold-start ignition and stable combustion at room temperature, with an average combustion chamber pressure of 0.93 MPa, an ignition delay time of 0.64 s, a pressure establishment time of 1.02 s, a characteristic velocity of 1168.7 m/s, and an average power of approximately 263 W in the decomposition zone circuit. Increasing the ignition voltage can shorten the ignition delay time (from 0.93 s to 0.47 s when increased from 60 V to 100 V), with 80 V identified as the optimal voltage overall. Although the arc has no significant effect on ignition response, it can effectively suppress low-frequency pressure oscillations. Reducing the electrode gap or optimizing the electrode orifice diameter to 0.8 mm can significantly improve ignition response characteristics. The study also reveals that the pressure oscillation frequency (<10 Hz) closely matches the current oscillation frequency, demonstrating that unstable propellant decomposition is the root cause of combustion instability. This research provides critical experimental evidence for the engineering design of catalyst-free ADN-based thrusters and offers important technical reference value for advancing the development of green high-performance space propulsion systems.
First, this paper focuses on the urgent demand for green space propulsion technologies and the inherent deficiencies of existing catalytic ignition approaches, and innovatively proposes an electrically ignited ADN-based thruster based on a combined scheme of resistive ignition and arc-assisted combustion. With the deepening concept of space sustainability, ADN-based liquid propellants, owing to their non-toxicity, high specific impulse, and favorable stability, have become the most promising green propellant alternative to hydrazine-based fuels. However, all ADN-based thrusters currently employed in engineering applications adopt the catalytic ignition technology route, which relies on highly active catalysts to achieve propellant decomposition and combustion, yet suffers from severe drawbacks: the catalyst cannot withstand temperatures exceeding 1500 K, the catalytic bed must be preheated to above 623 K prior to thruster startup, and insufficient preheating may lead to a "hard start" or even explosion. To overcome the technical bottleneck of catalytic ignition, this study for the first time designs an electric ignition experimental system as shown in Fig. 1, which mainly comprises the thruster, propellant supply system, ignition system, data acquisition system, and control system. The thruster adopts the structural design illustrated in Fig. 2, primarily consisting of a swirl injector, decomposition zone, combustion chamber, honeycomb multi-layer decomposition electrodes, arc electrodes, and a Laval nozzle, wherein the honeycomb electrode structure can enlarge the contact area between the propellant and the electrodes while suppressing secondary droplet splashing caused by micro-explosions. This electric ignition scheme requires neither catalyst nor preheating, and is expected to enable rapid cold-start of the thruster while avoiding the risk of hard start.
Second, the paper validates the feasibility of the electrically ignited thruster through systematic hot-fire tests, and investigates the effects of ignition voltage, arc loading time, electrode gap, and electrode orifice diameter on the ignition response and combustion characteristics of the thruster. Fig. 3 illustrates the thermal decomposition and combustion reaction pathways of the propellant during the hot-fire process. Under resistive heating, the propellant undergoes methanol dehydrogenation, water evaporation, and thermal decomposition of ammonium dinitramide, generating strongly oxidizing intermediates that subsequently undergo violent oxidation reactions with methanol and its dehydrogenation products in the combustion chamber, releasing substantial heat. Fig. 4 presents photographs of the thruster at four stages: pre-ignition, arc loading, ignition operation, and the end of the hot-fire test. It can be observed that during arc loading, the combustion chamber window exhibits a bright orange glow, and the light intensity further increases during the ignition operation stage, indicating that the propellant decomposition products are successfully ignited by the arc and achieve stable combustion. Under the operating conditions of 80 V ignition voltage, 3 mm electrode gap, and 0.8 mm electrode orifice diameter, the thruster achieves cold-start ignition at room temperature. The results of the 30-second hot-fire test, as shown in Fig. 5, demonstrate that the combustion chamber pressure is rapidly established after ignition, with an average pressure of 0.93 MPa, an ignition delay time of 0.64 s, a pressure establishment time of 1.02 s, and a characteristic velocity of 1168.7 m/s, which exceeds the design value of hydrogen peroxide thrusters of comparable thrust level. The voltage and current curves shown in Fig. 6 reveal that the average current in the decomposition zone circuit is 3.3 A, with an average power of approximately 263 W, and the resistance gradually increases and stabilizes as the propellant decomposition proceeds. The experiments also reveal periodic oscillations in the combustion chamber pressure, indicating the existence of combustion instability in the thruster.
Finally, the paper systematically analyzes the intrinsic correlations among ignition voltage, arc loading, electrode structure, and combustion instability, providing critical guidance for the optimal design of the thruster. The combustion chamber pressure curves and corresponding key performance parameters under different ignition voltages, arc loading times, electrode gaps, and electrode orifice diameters are respectively examined. The results indicate that increasing the ignition voltage can shorten the ignition delay time, with 80 V identified as the optimal voltage overall; although the arc is not a necessary condition for propellant ignition and combustion, it can significantly suppress pressure oscillations and improve combustion stability, while exerting no significant effect on ignition response characteristics; reducing the electrode gap can shorten both the ignition delay time and the pressure establishment time; when the electrode orifice diameter is increased from 0.3 mm to 0.8 mm, the average chamber pressure rises from 0.70 to 0.94 MPa and the ignition delay time decreases from 1.70 to 0.59 s, but further increasing it to 1.2 mm leads to performance degradation, as the excessively short residence time inhibits the decomposition reactions. As shown in Figs. 7 and 8, fast Fourier transform analysis reveals that the pressure oscillation frequencies are predominantly concentrated below 10 Hz, characteristic of low-frequency combustion instability, and the current oscillation frequency closely matches the pressure oscillation frequency with an opposite phase, confirming that unstable propellant decomposition is the root cause of combustion instability. Spray atomization characteristics analysis shows that the dominant frequency of droplet size fluctuations is above 50 Hz, indicating no direct coupling with pressure oscillations and only an indirect effect on the decomposition process. This electrically ignited thruster successfully overcomes the bottlenecks of catalyst activity degradation and explosion risk due to insufficient preheating inherent in catalytic ignition, offering advantages of extended lifespan and rapid startup. However, the energy consumption of approximately 263 W imposes higher demands on the spacecraft power system. This study provides critical experimental evidence and optimization directions for the engineering design of ADN-based thrusters.
Space: Science & Technology
Preliminary Hot-Fire Test of Ammonium Dinitramide-Based Thrusters Based on Electrical Ignition
2-Jul-2026