Assessment of Multimode Spacecraft Micropropulsion Systems

Berg, Steven P.; Rovey, Joshua L.

doi:10.2514/1.a33649

Cited by 35 publications

(6 citation statements)

References 16 publications

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“…A chamber temperature of 2,166 K is very high compared to other mono-propellant flame temperatures: the material employed for the chamber and thruster design is rhenium. It has a density of 20,800 kg/m 3 and a yield strength of 290 MPa [32].…”

Section: Thruster Designmentioning

confidence: 99%

Enabling Interplanetary Exploration for CubeSats with a Fully Chemical Propulsion System

Giordano¹

2023

JBIS

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Interplanetary CubeSat missions are currently becoming more popular, with a significant number of recently planned missions. The context of this paper is a Mars mission, starting from a parking orbit around Earth: the adoption of a chemical propulsion system for the Earth-Mars transfer phase is investigated, considering the recent technological developments for CubeSats. A trade-off of propulsion system type and propellant results in the choice of a mono-propellant system adopting the HAN-based propellant AF-M315E (ASCENT). The main challenge for the propulsion system is to fit inside a CubeSat standardised volume, which can range up to 24 U, for which the implementation of a suitable COTS micro-pump is considered. Finally, the complete architecture and design of the propulsion system is presented. This work demonstrates the feasibility of adopting full chemical propulsion for an interplanetary CubeSat mission, with consequent advantages in terms of transfer time and required power, at the cost of relatively small mass and volume left for the other subsystems. Even better results can be expected for interplanetary missions requiring slightly lower ΔV budgets, such as Near Earth Objects exploration or asteroid fly-by missions. Keywords: CubeSat, Chemical Propulsion, Green Mono-propellant, Interplanetary Mission, Mars Exploration

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Section: Thruster Designmentioning

confidence: 99%

Enabling Interplanetary Exploration for CubeSats with a Fully Chemical Propulsion System

Giordano¹

2023

JBIS

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“…The main benefit of this dual-mode propellant is to increase mission flexibility by sharing the same propellant between the high-thrust chemical thruster and the high-specific-impulse electric thruster equipped in one spacecraft. The satellite can be launched without a scheduled thrust history in advance, and both chemical and electrical propulsion systems can be activated according to in situ mission requirements. , A series of imidazole-based ionic liquid components, including 1-butyl-3-methylimidazolium nitrate ([Bmim][NO 3 ]), 1-butyl-3-methylimidazolium dicyanamide ([Bmim][DCA]), and 1-ethyl-3-methylimidazolium ethyl sulfate ([Emim][EtSO 4 ]), were assessed by Berg and Rovey . Among these ionic liquids, [Emim][EtSO 4 ] shows the best electrical propulsion performance, while the chemical propulsion performance is at the same level.…”

Section: Introductionmentioning

confidence: 99%

Catalytic Decomposition and Burning of a Dual-Mode Ionic Liquid Propellant

Yan²,

Wang

et al. 2021

Energy Fuels

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The ionic liquid dual-mode propellant, mainly composed of hydroxylammonium nitrate (HAN), 1-ethyl-3-methylimidazolium ethyl sulfate ([Emim][EtSO4]), and H2O, has shown good capabilities for both chemical and electrospray space propulsion. This work focuses on the chemical performance of the propellant by investigating its decomposition and burning processes. First, the thermogravimetric–Fourier transform infrared spectroscopy measurement is employed to identify different stages of the propellant decomposition, including water vaporization and decompositions of HAN and [Emim][EtSO4]. The effects of the Ir/Al2O3 catalyst, H2O mass fraction, and heating rates on the decomposition process are analyzed, and the gaseous products are measured. It is found that the decomposition of HAN and [Emim][EtSO4] is coupled with the oxidation of small intermediates. The kinetic parameters of overall catalytic decomposition/oxidation reactions at different stages are determined by fitting the thermogravimetric (TG) curves. Then, the burning of the propellant is organized in a well-designed optical-accessible catalytic bed with a controllable preheating device and a fixed flow rate of 2 mL/min. The ignition delay and burning evolution are determined from the time-resolved temperature sampling at four different axial positions. When the spatially averaged preheating temperature is 130–180 K, the ignition delay varies from 0 to 60 s, mainly controlled by the prior HAN decomposition rate. Finally, a one-dimensional model incorporating kinetic parameters from TG fitting is established to characterize the ignition process of the propellant, which quantitatively elucidates the effect of preheating on the ignition delay.

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“…Some relevant research has been carried out on the characteristics and performance of this propellant. Rovey's team from Missouri University of Science and Technology had conducted exploratory research on the application prospect of this ionic liquid in the dual-mode propulsion system, demonstrating the superiority and feasibility of this dualmode ionic liquid propulsion system [2][3][4]. Berg and Donius compared the performance of different kinds of ionic liquid propellants.…”

Section: Introductionmentioning

confidence: 99%

Experimental Study on Thermal and Catalytic Decomposition of a Dual-Mode Ionic Liquid Propellant

Gao¹,

Yao³

et al. 2021

E3S Web Conf.

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With the increasing of space activities and people’s awareness of environmental protection, it is necessary to develop a new non-toxic space propulsion system with high performance. Hydroxylammonium nitrate (HAN)/1-ethyl-3-methyl-imidazolium ethylsulfate([Emim][EtSO4]) blend ionic liquid propellant is a potential replacement with non-toxic and high-performance characteristics for hydrazine type liquid propellants, which can be used in both chemical and electrical propulsion system. This paper introduced the thermogravimetric experimental analysis (TGA-DSC) results of HAN/[Emim][EtSO4] ionic liquid propellant with the thermal decomposition and catalytic decomposition process. Its mass-loss process and exothermic process under different reaction conditions at a heating rate of 5K/min~15K/min were studied. Generally, the mass-loss results showed that there were four characteristic stages during the decomposition process of the HAN/[Emim][EtSO4] ionic liquids, which were the evaporation of the water solvent, decomposition of the HAN component, further decomposition of the [Emim][EtSO4], and slow loss of the residual substances. At the same time, two exothermic peaks were observed, which respectively corresponded to the decomposition of HAN and the further decomposition of [Emim][EtSO4]. Using catalyst can significantly reduce the decomposition temperature of the propellant and the residual mass. The contents in this paper proved that this propellant had a good application prospect within the catalytic ignition aerospace thruster.

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Assessment of Multimode Spacecraft Micropropulsion Systems

Cited by 35 publications

References 16 publications

Enabling Interplanetary Exploration for CubeSats with a Fully Chemical Propulsion System

Enabling Interplanetary Exploration for CubeSats with a Fully Chemical Propulsion System

Catalytic Decomposition and Burning of a Dual-Mode Ionic Liquid Propellant

Experimental Study on Thermal and Catalytic Decomposition of a Dual-Mode Ionic Liquid Propellant

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