Bimodal Propulsion System Using Ionic Liquid Propellant for Pico- and Nano-satellite Applications

Masuyama, Kento; Lozano, Paulo

doi:10.2514/6.2013-3962

“…System CE will be retained for the mission-defined analysis, but will not be included in the flexible-mission analysis since cold gas and electrospray are not compatible with the same propellant. 11 Finally, the helicon thruster is an electrodeless device, meaning there is no fundamental reason why any gaseous propellant could not be used. In the case of System MH, this will require an additional gas generator to decompose the AF315E prior to injection into the electric thruster.…”

Section: Multi-mode Propulsion Systemsmentioning

confidence: 99%

“…Given the specified parameters, and calculations from Eqs. (8)- (11), the remaining geometry of the divergence section, namely exit area and length are calculated through simple trigonometric relations. The thrust chamber geometry can be calculated through empirical means by Eqs.…”

Section: B Chemical Thruster Sizingmentioning

confidence: 99%

“…One includes the use of an ionic liquid propellant for chemical combustion or decomposition as well as for electrospray. 8,9,11 A specific propellant for this purpose is even under development. 12 This study will M examine this type of system, as well as others, specifically to compare these systems in reference to their multimode performance in reference to both mission-defined and flexible-mission scenarios.…”

Section: Introductionmentioning

confidence: 99%

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Assessment of Multimode Spacecraft Micropropulsion Systems

Berg

¹

,

Rovey

²

2017

Journal of Spacecraft and Rockets

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Multi-mode spacecraft micropropulsion systems which include a high-thrust chemical mode and high-specific impulse electric mode are assessed with specific reference to cubesatsized satellite applications. Both cold gas Freon-14 propellant and ionic liquid chemical monopropellant modes were investigated alongside pulsed plasma, electrospray, and helicon electric thruster modes. Systems involving chemical monopropellants have the highest payload mass fractions for a reference mission of a 500 m/s delta-V and 6U sized cubesat for electric propulsion usage below 55% of total delta-V. For higher electric propulsion usage, cold gas thrusters delivered a higher payload mass fraction due to lower system inert mass. Due to the combination of utilizing a common propellant for both propulsive modes, low inert mass, and high electric thrust, the cold-gas chemical/helicon-type electric combination had the highest mission flexibility, able to achieve a delta-V 10% lower than that of the largest delta-V system, but at roughly 500 days less burn time. A System utilizing a monopropellant thruster and electrospray thruster can achieve the largest delta-V, but with a burn time of over 600 days. This same system, however, can achieve the largest delta-V for missions requiring a thrust time of less than roughly 10 days.

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“…One includes the use of an ionic liquid propellant for chemical combustion or decomposition as well as for electrospray. 8,9,11 A specific propellant for this purpose is even under development. 12 This study will examine this type of system, as well as others, specifically to compare these systems in reference to their multimode performance in reference to both mission-defined and flexible-mission scenarios.…”

Section: Introductionmentioning

confidence: 99%

Assessment of Multi-Mode Spacecraft Micropropulsion Systems

Berg

¹

,

Rovey

²

2014

50th AIAA/ASME/SAE/ASEE Joint Propulsion Conference

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Multi-mode spacecraft micropropulsion systems which include a high-thrust chemical mode and high-specific impulse electric mode are assessed with specific reference to cubesatsized satellite applications. Both cold gas Freon-14 propellant and ionic liquid chemical monopropellant modes were investigated alongside pulsed plasma, electrospray, and helicon electric thruster modes. Systems involving chemical monopropellants have the highest payload mass fractions for a reference mission of a 500 m/s delta-V and 6U sized cubesat for electric propulsion usage below 55% of total delta-V. For higher electric propulsion usage, cold gas thrusters delivered a higher payload mass fraction due to lower system inert mass. Due to the combination of utilizing a common propellant for both propulsive modes, low inert mass, and high electric thrust, the cold-gas chemical/helicon-type electric combination had the highest mission flexibility, able to achieve a delta-V 10% lower than that of the largest delta-V system, but at roughly 500 days less burn time. A System utilizing a monopropellant thruster and electrospray thruster can achieve the largest delta-V, but with a burn time of over 600 days. This same system, however, can achieve the largest delta-V for missions requiring a thrust time of less than roughly 10 days. NomenclatureA c = combustion chamber cross sectional area, [m 2 ] A t = throat area, [m 2 ] C F = thrust coefficient C = effective exhaust velocity, [m/s] D c = combustion chamber diameter, [m] D t = throat diameter, [m] EP = electric propulsion usage fraction F = thrust, [N] F tu = ultimate strength of material, [N/m 2 ] f inert = inert mass fraction g 0 = acceleration of gravity, [m/s 2 ] I sp = specific impulse, [s] I sp,chem = chemical mode specific impulse, [s] I sp,elec = electric mode specific impulse, [s] I sp,mm = multi-mode effective specific impulse, [s] L c = combustion chamber length, [m] L * = characteristic combustion chamber length m 0 = initial mass of spacecraft, [kg] m c = combustion chamber mass, [kg] m chem = mass of chemical propellant, [kg] m elec = mass of electric propellant, [kg] m f = final mass of spacecraft, [kg] m f1 = mass of spacecraft after first burn, [kg] 2 m inert = inert mass, [kg] m pay = payload mass, [kg] m PPU = mass of power processing unit, [kg] m prop = propellant mass, [kg] m sa = mass of solar array, [kg] m tank = mass of propellant tank, [kg] P b = burst pressure, [Pa] P c = chamber pressure, [psi] P e = nozzle exit pressure, [Pa] P thr = electric thruster power, [kW] r c = combustion chamber radius, [m] r t = throat radius, [m] t b = thruster burn time [day] t w = wall thickness, [m] α = nozzle divergence half-cone angle, [degrees] ΔV = velocity increment, [m/s] ε = nozzle expansion ratio η t = thrust efficiency θ c = convergent section angle, [degrees] γ = specific heat ratio λ = nozzle divergence correction factor φ tank = empirical tank sizing parameter ρ prop = propellant density, [kg/m 3 ] ρ w = wall material density, [kg/m 3 ]

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