Low Specific Impulse Research Articles

Combustion behavior of NMP-001, a nitromethane-based green rocket propellant

Hydrazine’s carcinogenic and toxic nature makes its use in space propulsion expensive and necessitates careful launch campaign planning, which may cause delays. One already established alternative is to use energetic salt-based ionic solution propellants. However, the costs remain high because these salts are costly and highly regulated energetic materials. Additionally, propellants based on these energetic salts produce much higher temperatures during decomposition and combustion, as is the case for hydrazine. Consequently, expensive refractory metal-based alloys and high-temperature oxidation-resistant catalytic reactors must be used, which further drives up the costs. Alternatively, hydrogen peroxide may be used, though only in applications with a low specific impulse requirement. This study explores an alternative approach: using nitromethane, a commonly available laboratory solvent. NMP-001 is a nitromethane-based monopropellant that contains additives that allow for low-pressure combustion and reduce sensitivity to mechanical stimuli. Multiple combustion chamber configurations with varying characteristic lengths were tested in rocket combustion experiments. With the 11.5 m L∗-configuration, a 15.1 bar low-pressure limit could be determined. The 7.3 and 5.0 m L∗-configurations yielded a 16.2 and 28.3 bar low-pressure combustion limits, respectively. These outcomes are significantly below the 30 bar limit reported in the literature. Despite using only a simple heat-sink combustion chamber with a non-sophisticated injector element and a non-optimized propellant formulation, combustion was 83 to 95 % efficient with a C* ranged between 1133 to 1302 m/s. These figures are higher than for hydrogen peroxide and are comparable to the theoretical performance of state-of-the-art green liquid monopropellants.

Monoethanolamine-Based Fuels Hypergolic with Rocket-Grade Hydrogen Peroxide

The object of this work is to study the ignition performances of promising catalytically promoted amino-based green fuels to be used in combination with rocket-grade hydrogen peroxide. The main hypergolic parameter, the ignition delay time, was experimentally determined with an automated drop test setup. Additionally, the kinematic viscosity was experimentally measured, while the CEA2 software was used to determine the ideal rocket performances of the propellants. Three inorganic copper salts were selected as catalysts to be used in combination with monoethanolamine in concentrations ranging from 0.5 wt% to 20 wt%. Then, N,N-dimethylethylenediamine was introduced as part of a blend with monoethanolamine to target the high viscosity and low gravimetric specific impulse of the fuel for the pure monoethanolamine case. Due to the reduced monoethanolamine and low additive content, some formulations were observed to be characterized by gravimetric specific impulse higher than 320 s, and kinematic viscosity lower than 5 cSt while retaining ignition delay times shorter than 30 ms with 98 wt% HTP. Finally, the addition of ethanol to the blend was preliminarily investigated to improve the additive solubility. The effects on the ignition delay time were found to depend on the concentration ratio of the two amines.

Novel Hypergolic Green Fuels with Hydrogen Peroxide for Propulsion Systems

By combining the advantages of chemicals from two different classes, a series of catalytically promoted green hypergolic fuels named polyamine/alkanolamine-based hypergolics (PAHyp) with highly concentrated hydrogen peroxide (96%) as an oxidizer was developed. In this paper, a novel recipe based on N,N,N′N′-tetramethylethylenediamine (TMEDA) and N-methyldiethanolamine (MDEA), named PAHyp 1, is characterized. Samples with different volume proportions of TMEDA and MDEA catalyzed with 0.5–4 wt% of copper salts were prepared. It was demonstrated that, by adding low catalyst loadings of 1–2 wt%, ultrafast ignition (as low as 8 ms) can be measured. Fast ignition is important to avoid a hard start in the startup phase, and reduced catalyst loading is important to avoid loss of performance in terms of specific impulse. To evaluate the system’s performance, an orbital transfer maneuver of a geostationary satellite that burns monomethylhydrazine with nitrogen tetroxide was considered. It was demonstrated that when using green formulations based on TMEDA/MDEA, smaller propellant tanks are required (because of the higher density of the green propellants), though more propellant mass is required due to a slightly low specific impulse. Besides good performance and reliable ignition, good storability is crucial for long-term space applications. Remarkably, although TMEDA is sensitive to air, visual inspection, ignition tests, and Fourier-transformed infrared spectroscopy analysis revealed that formulations with volume proportions of at least 60% of MDEA stored in vials filled with air survived with no signs of degradation after 14 months. Finally, to meet the requirement of long-term missions of several years in space, a hypergolic tripropellant feed system was proposed.

Research progress on the catalytic and thermal decomposition of ammonium dinitramide (ADN).

Ammonium dinitramide (NH4N(NO3)2, ADN) is regarded as a promising oxidizer due to its low signature and high specific impulse. Generally, ADN undergoes exothermic decomposition above 140 °C accompanied by the byproduct of ammonium nitrate (AN). The inevitable endothermic decomposition of AN decreases the overall heat release, and so there is a need to develop efficient catalysts to guide ADN decomposition along desired pathways with a lower decomposition temperature and higher heat release. A suitable catalyst should be able to withstand the harsh conditions in a thruster to achieve a stable thrust force, which poses a huge obstacle for manufacturing a stable and active catalyst. This review gives a comprehensive summary of the thermal and catalytic decomposition pathways of ADN for the first time, which is expected to deepen the understanding of its reaction mechanism and provide useful guidance for designing prospective catalysts toward efficient ADN decomposition.

Directional combustion induced by 3D aligned carbon channel arrays for combustion chamber-free micro-thruster

Micro-solid propellant thrusters are widely used in propulsion and attitude adjustment of micro-spacecraft, and nano-satellites prospect. However, they suffer from low specific impulse owing to the existence of the conventional structural parts (combustion chamber and nozzle) that account for a large weight ratio. Herein, we report a combustion chamber-free micro-thruster consisting of balsa wood-derived carbon aerogel as well as ammonium perchlorate (AP) and Ti3C2 MXene supported on carbon aerogel. The microstructure engineering design of carbon aerogel give it 3D aligned channel arrays structure and defect rich carbon framework. The 3D aligned channel could guide the burning fluid in the channel to form directional flame, thus converting heat energy into kinetic energy with absence of the additional combustion chamber and nozzle. The defect-rich carbon aerogel with high reactivity could participate in the combustion reaction as fuel. To further promote the combustion performance of the chamber-free micro-thruster, Ti3C2 MXene is applied as reactive combustion catalyst to accelerate the thermal decomposition of AP. Coupled with the advantageous of chamber-free, high combustion efficiency, and efficient heat and mass transfer in the channel structure, the Ti/CA/AP composite can be directly applied as a micro-thruster and generates a high specific impulse of 2044.0 N·s/kg.

Alternative Propellant Nuclear Thermal Propulsion Engine Architectures

Recently, NASA’s Artemis Program has considered sustainably returning humans to the moon with in situ resource utilization as the central focus and considers nuclear thermal propulsion (NTP) to be a viable propulsion system candidate for deep space missions. This program also considers the moon to function as a staging area for deep space missions due to its strategic orbit around Earth. The Lunar Crater Observation and Sensing Satellite discovered the abundancy of water, ammonia, and other volatiles in permanently shadowed regions on the moon. Because NTP can theoretically use any fluid as a propellant, water and ammonia are two potential alternative propellants to hydrogen that do not require postprocessing to be used directly in situ. Alternative propellant NTP engine models were developed to examine the performance of water and ammonia. It was found that bleed cycle architectures are infeasible with current turbine materials, but expander cycles have acceptable turbine inlet temperatures with marginal performance differences from bleed cycles. At reactor power levels comparable to hydrogen NTP engines, water doubles the thrust while ammonia increases it by 70%, although at a much lower specific impulse. The largest contributor to the life of the engine was the maximum temperature of the SiC cladding for water and the maximum temperature of the fuel for ammonia where higher values resulted in shorter engine life.

Application and performance estimation of Mg/CO2 engine on Mars

The production of return propellant through Martian in-situ resources has been proposed in the assumptions of Mars exploration missions. The Mg/CO2 scheme is one of these approaches in which powdered magnesium comes from the Earth as fuel and CO2 is collected on the surface of Mars as the oxidant. The propellant of Mg/CO2 engines has a good storage stability, has no chemical reactions in the in-situ resource utilization process, and contains a simple structure; thus, it is regarded as one of the technologically less complex choices for missions on Mars. However, the low specific impulse of Mg/CO2 engines requires evaluating their applications and performance. Three Mg/CO2 application scenarios (Ascent Vehicle, Hopper, and Two-stage Returnable Aircraft) are reviewed here, and two of them are estimated in detail from the established motion models. It is found that when a Mg/CO2 engine is used as an Ascent Vehicle, a two-stage rocket is more suitable. The propellant carried from Earth for Mg/CO2 engines could be reduced by 27.7% compared with the conventional storable UDMH/N2O4 (specific impulse is 330 s) two-stage rocket. The saved propellant is largest when the total velocity pulse is approximately 3430 m/s. When Mg/CO2 engines are used as a jump vehicle “Hopper”, the increased maximum thrust, engine working time, and total jump times decrease the marginal benefit, which means that these parameters cannot be too large. Based on the Ascent Vehicle and Hopper schemes, a Two-stage Returnable Aircraft is proposed for material and personnel circulation between different Mars bases.

Experimental study of a porous electrospray thruster with different number of emitter-strips

The electrospray thruster is becoming popular in space propulsion due to its low power and high specific impulse. Before this work, an electrospray thruster based on a porous emitter was developed. In order to achieve larger and more stable thrust, the thruster was redesigned, and the influence of the space between strips on thrust was studied. Four types of emitter were tested, and they had 1, 3, 4 and 14 emitter-strips on the emission surface of the same size respectively. According to the experimental results, the maximum extraction voltage and emission current of the four thrusters are different under stable operational conditions. The measured stable emission currents and extraction voltages were −500 μA/−5000 V, −1570 μA/−3800 V, −1200 μA/−3800 V, and −650 μA/−4500 V, respectively. Increasing the number of strips may not result in the emission current increasing, but changing the stable operational range of the emission current per strip and the extraction voltage. The maximum stable operational extraction voltages of 3 and 4 emitter-strips are lower than those of 1 and 14 emitter-strips, but the emission currents are higher than those of 1 and 14 emitter-strips. Time-of-flight mass spectrometry was used to analyze the mass distribution and obtain the performance of the thruster in the case of thrusters with 1 and 3 emitter-strips. Both of their plumes were composed of very small ion cluster (the pure-ion regime), and their thrusts were 80.1 μN, 219.2 μN with specific impulses of 5774 s, 5047 s, respectively.

Revolutionizing Spaceflight: A Study on Electric Propulsion and Air-Breathing MPDs

Modern liquid-fuel rocket propulsion harbors a number of great limitations. Among those is the weight of fuel, which makes up more than 90% of the mass of the SpaceX Falcon 9 (NASA, 2018). Electric propulsion has been used for decades as an alternative to liquid-fuel rockets due to low propellant requirements and high specific impulse. Although electric thrusters have strictly been used in non-atmospheric conditions, recent innovations attempt to expand its use to airspace. This quasi-experimental study focuses on the creation of an air-breathing magnetoplasmadynamic (MPD) thruster, with attempts being made to maximize the efficiency of the engine. Immense safety concerns prevented testing from occurring after the engine was built. However, the estimated performance of the built MPD is compared to a multitude of existing forms of electric propulsion, from Hall-effect thrusters to electrodynamic tethers. The concluding evidence suggests that air-breathing MPDs are not currently viable, high-power photon thrusters being of greater use in atmospheric conditions. Further research focusing on decreasing atmospheric breakdown voltage and increasing mirror reflectance of photon thrusters is suggested.

From heart drug to propellant fuels: Designing nitroglycerin-ionic liquid composite as green high-energy hypergolic fluids

New-generation green bipropellants based on ionic liquids for aircraft have attracted the attention of researchers. However, the major defects of ionic liquid propellants are their low specific impulse and incomplete combustion, which limits their practical application. In this work, a series of novel hypergolic fluids were prepared by mixing oxygen-rich nitroglycerin (NG) with fuel-rich dicyanamide-based ionic liquids. The complete dissolution of NG in ionic liquids makes these hypergolic fluids became self-supplying oxygen composite fuels, which generate much less solid residual products than pure hypergolic ionic liquids (HILs). Moreover, the addition of NG inhibited the micro-explosion between fuel and white fuming nitric acid (WFNA), extended the contact time between fuel and oxidizer. And a proper amount of NG in the fuel could suppress the secondary combustion. These composite fuels have high density (>1 g cm−3), low viscosity (as low as 13.49 cP), wide liquid operating ranges and acceptable ignition delay time. Additionally, the specific impulses of the composite fuels were higher than pure HILs. For the first time, this work provides a simple and reliable method to control the ignition properties of HIL by adding oxygen-enriched additives.

Miniaturized rotating magnetic field–driven plasma system: proof-of-concept experiments

The development of plasma-based propulsion thrusters for spacecraft has seen a rapid growth over the past few decades, with the number of spacecraft including small satellites and Cubesats increasing exponentially. Although traditional chemical propulsion is still widely employed in space flights, it cannot meet the more challenging requirements for deep space travel due to low specific impulse. Electric propulsion thrusters have already helped humans travel further from Earth and have the potential to be developed for interstellar flights due to their advantages such as high velocity increments, long operational lifetimes, high impulse-to-weight ratios and high impulse-to-power ratios. The electrodynamic thrusters have significant potential for applications in the remote regions of space, and several types of electrodynamic plasma thrusters are currently under investigation. In this paper we present conceptual experiments to study a miniaturized Rotamak-type device initially proposed for the controlled thermonuclear fusion, with a view to assess its potential for the application as a small space thruster. An outline of the physical characteristics of the experiments that has been carried out, and measurements were done to try to elucidate the important mechanisms at work in the Rotamak, which will help design next-generation thruster capabilities. A discussion is also presented about the Rotamak systems and the opportunities they present for space applications.

Ionic liquid electrospray behavior in a hybrid emitter electrospray thruster

Electrospray thrusters are most suitable for micro-nano satellites in advantage of small size, low power consumption and high specific impulse. However, conventional electrospray thrusters with capillary emitters operate at single mode with limited range of thrust. In order to expand range of thrust, novel electrospray thrusters with hybrid emitters which are capable of operating at dual-modes are recommended. Molecular dynamics simulation is used to study the behaviors of the novel electrospray thruster in this work, using EMIM-Tf2N as propellant. Electrospray behaviors for the capillary emitter and the hybrid emitter are compared at the same operating conditions, showing that droplets are more easily emitted from the capillary emitter. For the hybrid emitter, the externally-wetted emission mode is displayed at a low flow rate. Only ions are emitted from the hybrid emitter because of the stronger electric field and larger hydraulic impedance. Meanwhile the change in range of mean diameter and charge to mass ratio of particles are relatively small as compared to the capillary emitter. But the hybrid emitter transits into a capillary emission mode at a high flow rate. Furthermore, analysis of particle charge shows that if there exist particles with charge below the Rayleigh limit, droplets can be produced in electrospray beam.

Unintended consequences with laser nudging or re-entry of satellites

We measured the impulse coupling coefficient Cm (target momentum per joule of the incident laser light) and ejecta velocity vE with glass-covered Si solar (photovoltaic) cells following the irradiation by repetitive 71 ps, 1064 nm laser pulses at fluence of order 10 J/cm2 in vacuum. We measured high Cm values up to 600 N/MW, but very low specific impulse Isp = vE/go of order 10 s. Thousands of glass particles 20–1500 μm in size were ejected and found on the floor of the vacuum chamber close to the target. At lower fluences, larger pieces were ejected. While pulsed laser coupling data exist on typical metals used for satellite construction, to our knowledge, this is the first data on solar arrays which comprise a large fraction of the total area exposed to a laser pulse intended to nudge or re-enter a satellite. On bare Si, Cm was about two orders of magnitude lower, and no large pieces were dislodged. The consequence of these data are that lasers for nudging or re-entering defunct satellites should be focused carefully on the metal parts, avoiding solar arrays, to avoid creating a cloud of microscopic debris, enhancing rather than ameliorating the space debris problem. Further, data on MLI (multi-layer insulation) and other complex layered structures should be obtained to guarantee against other unintended consequences.

Influence of propellant leakage from pump area into turbine area on turbo-pump operation stability

There is an increasing trend to liquid-propellant rocket engines which run on eco-friendly storable propellant. This trend is mostly dictated by the refusal to use traditional toxic storable propellant in many countries. The most widespread eco-friendly storable propellant is hydrogen peroxide with kerosene. Though, this propellant has a lower specific impulse in comparison with traditional liquid oxygen with kerosene. To compensate the loss of specific impulse, there is a reason to design a staged combustion engine. Evidently, the turbopump is the most complicated system in the staged combustion propulsion system. This fact makes research devoted to turbo-pumps a top priority. The paper aims to determine the influence of propellant leakage from the pump area into the turbine area and create recommendations which would allow organizing the stable operation of turbopump. As a result of turbopump staged combustion cycle testing, a conclusion had been made that leakage, which opens during the test, significantly influences the stability of turbopump operation. Depending on the amount of leakage, the turbine generated power drop was between 20 and 45%, which led to a decrease in rotation speed and outlet pressure of the pump. During the R&D process, a way of leakage influence elimination had been offered. Formulated recommendations may be used during the design process of the turbopump for staged combustion liquid propulsion systems.

Direct Force Measurement of Pulsed Laser Ablation of Asteroid Simulants

Subnanosecond laser ablation of asteroid simulants has been studied, and direct force measurements were made. A 0.7 ns pulsed laser was used, with a 40 kHz pulse repetition frequency, 1064 nm wavelength, and 33 W average power. High-fidelity simulants for carbonaceous chondrite meteorites of types CM, CR, and CI, as well as pyroxene rock and 6061 aluminum were ablated. The force response over 500 ms of ablation was measured using a load cell. In addition, the samples were weighed before and after ablation to determine the mass removal rate. Based on these measurements, the momentum coupling coefficient, specific impulse, and ablation efficiency were calculated. The asteroid analog materials were found to have a modestly higher momentum coupling coefficient than aluminum, although the specific impulse and efficiency were much lower. For asteroid mitigation missions, the asteroid itself is used as the propellant, and so a low specific impulse poses no problem.

Influence of screen gird aperture diameter in outer region on performance of dual-mode ion thruster

To eliminate the erosion in the apertures at the edge of the decelerate grid of a dual-mode ion thruster three-grid optical system, a new method of reducing the diameter of the apertures in the outer region of the screen grid is proposed. In order to investigate the method of influencing the discharge loss and the uniformity of the beam current density and also reducing the aperture erosion region and erosion rate of the decelerate grid, two kinds of three-grid optical systems are designed and fabricated with the same material and physical parameters except the diameter of the apertures in the outer region of the screen grid. The first kind is that the diameter of the aperture in the outer region of the screen grid is equal to that in middle region of the screen grid, it is designated as the same aperture diameter grid optical system. The second kind of three-grid optical system is that the diameter of screen apertures whose center distance from the center of grid is larger than 0.95 times the beam radius, is reduced by 26% and therefore the physical transparency of optical system is reduced by 8.8%, it is designated as the small aperture diameter grid optical system. The comparison between the same aperture diameter grid optical system and small aperture diameter grid optical system is performed by assembling them into a 30-cm-diameter dual-mode ion thruster, and the ion thruster performance test, beam flatness test and 600-h-endurance test are conducted in the two typical operation modes, namely the low thrust-high specific impulse mode and large thrust-high power mode. A comparison of small aperture diameter grid optical system with the same aperture diameter grid optical system shows that the discharge loss of the 30-cm dual-mode ion thruster in the low thrust-high specific impulse mode and large thrust-high power mode are reduced by 10% and 21% respectively, the beam flatness is decreased by 3% and 10% respectively, the number of rows of the apertures which are sputtered by beam ions at the edge of the decelerate grid is reduced from 5 to 1. In addition, the erosion rate is significantly reduced and the erosion phenomenon disappears after 900-h-long-duration operation. These results signify that reducing the screen grid aperture diameter in the outer region is an effective method to eliminate aperture erosion at the edge of the decelerate grid of a dual-mode ion thruster three-grid optical system and that this method will not reduce the efficiency of the thruster, but cause the beam current uniformity to worsen.

Thermodynamics analysis of a turbojet engine integrated with a fuel cell and steam injection for high-speed flight

The high-speed flight can be achieved by high-performance aero-engines. A traditional turbojet engine with steam injection is hardly operated at Mach 3.5–5 owing to the low specific impulse. The performance of turbine-less jet engines is outstanding. No turbines exist in the engine and compressors are powered by fuel cells. Therefore, the limitation of the highest combustion temperature caused by turbines disappears. In addition, fuel cell exhaust can still be used to burn and expand to output propulsion power. In this paper, the turbine-less jet engines integrated with a SOFC and steam injection are proposed to achieve high-speed flight. In order to analyze the performance of the engine, the thermodynamic model is built. Effects of thermodynamic parameters such as water-air ratio are studied. The main results are as follows: the performance of the turbine-less engine is significantly improved by integrating with steam injection. The augment of specific impulse and specific thrust of the engine is 13.13% and 51% respectively compared with the pre-cooling turbojet engine at the water-air ratio of 0.013. The specific impulse of the engine is 1.15 times that of the pre-cooling turbojet engine at Mach 4. Meanwhile, the specific thrust is increased by 26%.

Miniature ion thruster ring-cusp discharge performance and behavior

Miniature ion thrusters are an attractive option for a wide range of space missions due to their low power levels and high specific impulse. Thrusters using ring-cusp plasma discharges promise the highest performance, but are still limited by the challenges of efficiently maintaining a plasma discharge at such small scales (typically 1–3 cm diameter). This effort significantly advances the understanding of miniature-scale plasma discharges by comparing the performance and xenon plasma confinement behavior for 3-ring, 4-ring, and 5-ring cusp by using the 3 cm Miniature Xenon Ion thruster as a modifiable platform. By measuring and comparing the plasma and electron energy distribution maps throughout the discharge, we find that miniature ring-cusp plasma behavior is dominated by the high magnetic fields from the cusps; this can lead to high loss rates of high-energy primary electrons to the anode walls. However, the primary electron confinement was shown to considerably improve by imposing an axial magnetic field or by using cathode terminating cusps, which led to increases in the discharge efficiency of up to 50%. Even though these design modifications still present some challenges, they show promise to bypassing what were previously seen as inherent limitations to ring-cusp discharge efficiency at miniature scales.

Chip based MEMS Ion Thruster to significantly enhance Cold Gas Thruster Lifetime for LISA

Micropropulsion is a key component for ultraprecise attitude and orbit control required by the eLISA mission. LISA pathfinder uses cold gas micro thrusters that are accurate but require large tanks due to their very low specific impulse, which in turn limits the possible mission duration of the follow up eLISA mission. Recently, we developed a compact MEMS ion thruster on the chip with a size of only 1cm2 that can be simply attached to a gas feeding line like the one used for cold gas thrusters. It provides a specific impulse greater than 1000 s and only requires a single DC voltage. Since the operating principle is based on field emission, very low thrust noises similar to FEEP thrusters are expected but with gas propellants. The MEMS ion thruster chip could be mounted in parallel to the existing gold gas system providing high Isp and therefore long mission durations while leaving the cold gas system in place. To enable a possible mission extension, the MEMS ion thruster could take over from the cold gas system as a backup while maintaining the existing micropropulsion thruster system with its heritage therefore minimum risk.

5kW級ホールスラスタRAIJINの性能解析

A Hall thruster system for in-space propulsion has been developed as a part of a Japanese collaborative research and development project aiming at practical use of a high power anode layer type Hall thruster. Its thrust was measured at the ISAS/JAXA vacuum chamber facilities using a dual-pendulum thrust stand developed at the University of Tokyo. In the low specific impulse mode operation, thrust efficiency of 64% and thrust to power ratio of 80mN/kW were achieved. In the high specific impulse mode operation, the thruster was successfully operated at the acceleration voltage of up to 1,000V.