Assessment of Imidazole-Based Ionic Liquids as Dual-Mode Spacecraft Propellants

48th AIAA/ASME/SAE/ASEE Joint Propulsion Conference &Amp;amp; Exhibit

2012

The ionic liquid [Bmim][DCA] is a propellant candidate in a standalone electrospray thruster or in a dual-mode propulsion system consisting of a chemical system and an electrospray system. Since limited published data exists for [Bmim][DCA], the electrospray characteristics are relatively unknown. Emission testing of the ionic liquid has been conducted to characterize the [Bmim][DCA] electrospray plume for both an external flow titanium needle and internal flow capillary. Mass spectrometric, retarding potential, and angle distribution measurements were collected for the positive polarity ions emitted from [Bmim][DCA] wetted emitters with nominal extraction voltages between ~1 kV to ~2.5 kV. The titanium needle operated at a sizably reduced liquid flow rate in comparison to the capillary. As such, only the major species of Bmim + ([Bmim][DCA]) n with n=0,1 were identifiable in the quadrupole measurement range of 0-1000 amu and were formed at or near the needle potential. A typical needle angle distribution was found in these measurements. For the capillary emitter, flow rates from 0.27 nL/s to 2.18 nL/s were used to investigate corresponding alterations in the electrospray beam. The aim of the investigation was to ascertain the ability to "tune" or "dial-in" an electrospray thruster to specific ion or droplet sizes and thus specific performance levels. Unlike the limited species observed from the needle emission, the capillary measurements indicated the presence of n=0,1,2,3,4 cation species with large mass droplet contributions. The lowest flow rates indicated the highest levels of ions in the measurement range of 0-1000 amu with a mix of large mass droplets. For increasing flow rate, species < 500 amu ceased to exist leaving only the n=2,3,4 species mixed with large mass droplets in the electrospray beam. All ion species exceeded the quadrupole mass range at the upper flow rates. Ions emitted from the capillary were formed at levels below the emitter potential. Ohmic losses in the ionic liquid are likely the cause for the less energetic ions. Angular distribution measurements indicated broadening of the beam current and mass distribution for increasing flow rates. Nomenclature D c = transport capillary inner diameter (µm) D n = capillary needle emitter inner diameter (µm) K = ionic liquid electrical conductivity (S/m) L c = transport capillary length (cm) L n = capillary needle emitter length (cm) P 0 = reservoir pressure (Torr) P VC = vacuum chamber pressure (Torr) 2 P n = capillary needle emitter pressure (Torr) q/m = specific charge (C/kg) Q = ionic liquid volumetric flow rate (nL/s) V N = emitter voltage (V) V Ext = extractor plate voltage (V) β n = capillary length/diameter ratio coefficient γ = ionic liquid surface tension (N/m) ε = ionic liquid dielectric constant µ l = ionic liquid viscosity (cP) ρ = ionic liquid density (kg/m 3 )

Section: Introductionmentioning

confidence: 99%

Section: Introductionmentioning

confidence: 99%

Capillary Extraction of the Ionic Liquid [Bmim][DCA] for Variable Flow Rate Operations

Miller¹,

Prince²,

48th AIAA/ASME/SAE/ASEE Joint Propulsion Conference &Amp;amp; Exhibit

2012

“…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

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

2014

Self Cite

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 ]

“…[3][4][5][6][7][8] For example, use of a separate chemical rocket to escape earth gravity avoids a long spiral trajectory characteristic of an electric burn, while a high-specific impulse electric burn in interplanetary space saves propellant mass over a chemical rocket. 7 However, it has been shown that even greater mass savings can potentially be realized through the use of shared propellants [9][10][11] or shared hardware. 12 The overidding question, however, when assessing multi-mode propulsion system performance is whether inherent thruster performance losses accrued as a compromise in integrating either propellant or hardware outweigh the anticipated benefits.…”

Section: Introductionmentioning

confidence: 99%

Performance Analysis of an Integrated Multi-Mode Chemical Monopropellant Inductive Plasma Thruster

Berg

49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference

2013

Self Cite

A novel multi-mode spacecraft propulsion concept is presented. The concept combines chemical monopropellant and electric pulsed inductive thruster technology to include shared propellant and shared conical nozzle. Geometry calculations show that existing conical pulsed inductive thruster experiments are typical of large (1000-4000 N) chemical monopropellant thruster nozzles. Performance and propulsion system mass required to accomplish a 1500 m/s delta-V with a 500 kg payload was calculated for geometries including 20-55 degree divergence angles. Results show that combining nozzle geometry is not beneficial in terms of propulsion system mass for small nozzle divergence angles, however using a nozzle with a 55 degree divergence angle results in a 1-2% reduction in propulsion system mass compared to an equivalent thrust system utilizing a separate chemical bell nozzle and flat coil PIT device despite having 19% lower chemical specific impulse and 18% lower electric thrust efficiency. Results suggest that using even larger divergence angles could yield even more benefit. Nomenclature= current sheet mass distribution function f inert = inert mass fraction g 0 = acceleration of gravity I 1 = powertrain currentcombustion chamber length L * = characteristic combustion chamber length, or inductance ratio M = mutual inductance m = current sheet mass m bit = mass bit m c = combustion chamber mass m cables = mass of electrical cables 2 m inert = inert mass m pay = payload mass m PPU = mass of power processing unit m prop = propellant mass m sa = mass of solar array N = PIT nozzle angle scaling parameter P b= burst pressure P c = chamber pressure P e = nozzle exit pressure R e = external circuit resistance R p = plasma resistance r = radial position of current sheet r c = combustion chamber radius r coil.= coil radius r t = throat radius t = time t w = wall thickness V = voltage v r = radial current sheet velocity v z = axial current sheet velocity z = axial position of current sheet α = nozzle divergence half-cone angle α r = radial dynamic impedance parameter α z = axial dynamic impedance parameter ΔV = velocity increment ε = nozzle expansion ratio η t = thrust efficiency θ c = convergent section angle γ = specific heat ratio λ = nozzle divergence correction factor φ tank = empirical tank sizing parameter ψ 1 = powertrain critical resistance ratio ψ 2 = plasma critical resistance ratio ρ = current sheet density ρ prop = propellant density ρ w = wall material density superscript * = nondimensionalized quantity subscript 0 = initial quantity