Thermal nonequilibrium in a low-power arcjet nozzle

Zube, Dieter M.; Myers, Roger

doi:10.2514/3.23657

Cited by 31 publications

(26 citation statements)

References 10 publications

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“…In addition, the rarefied gas effect is not considered in the modeling. The N 2 vibrational and rotational temperatures measured by Zube and Myers [5] are shown in Fig. 11b for Nozzle II with simulated hydrazine as propellant, mass flow rate of 47.6 mg/s, and arc current of 9 A.…”

Section: Resultsmentioning

confidence: 99%

“…Numerical simulation of the NASA 1-kW class arcjet thruster with hydrogen or hydrazine as the propellant was also performed, and the predicted results for the variation of gas temperature and axial velocity along the nozzle axis and for the radial distributions of the axial velocity and temperature at the thruster exit agree reasonably well with the corresponding experimental results. Comparison of predicted gas temperature variation along nozzle axis with the experimental results of Zube et al [5] (a) (b) Fig. 11 Comparisons of the computed results and the experimental data concerning the radial profiles of the axial velocity at the exit plane (a) and the axial variations of gas temperature along the nozzle axis (b), for Nozzle II with simulated hydrazine as the propellant.…”

Section: Discussionmentioning

confidence: 96%

“…The all-speed SIMPLE algorithm [24], which is incorporated into the modified FAST-2D program to simulate the subsonic-supersonic flow, is used to solve the governing equations (1)(2)(3)(4)(5), associated with the auxiliary relations (6-9) and (11) and the specified boundary conditions, to obtain the distributions of the velocity components, pressure and specific enthalpy (or temperature) within the whole thruster nozzle. Altogether 89 (z-direction) 9 30 (r-direction) grid points are employed in this study, and a special numerical test shows that almost mesh-independent results have been obtained.…”

Section: Modeling Approachmentioning

confidence: 99%

“…different specific impulses, different thrust efficiencies, different arc voltages, etc.) are reported [2][3][4][5][6][7][8][9][10][11][12][13][14][15][16][17][18][19] and those results are often obtained for different power levels, different thruster construction or sizes and for different thruster operating parameters. In order to better clarify the effect of the propellant composition on the thruster characteristics, our modeling studies are performed using different kinds of the propellants but for the same thruster construction/sizes and the same operating conditions.…”

Section: Introductionmentioning

confidence: 99%

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Modeling Study on the Flow, Heat Transfer and Energy Conversion Characteristics of Low-Power Arc-Heated Hydrogen/Nitrogen Thrusters

Wang

Geng

Chen

et al. 2010

Plasma Chem Plasma Process

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A modeling study is conducted to investigate the effect of hydrogen content in propellants on the plasma flow, heat transfer and energy conversion characteristics of lowpower (kW class) arc-heated hydrogen/nitrogen thrusters (arcjets). 1:0 (pure hydrogen), 3:1 (to simulate decomposed ammonia), 2:1 (to simulate decomposed hydrazine) and 0:1 (pure nitrogen) hydrogen/nitrogen mixtures are chosen as the propellants. Both the gas flow region inside the thruster nozzle and the anode-nozzle wall are included in the computational domain in order to better treat the conjugate heat transfer between the gas flow region and the solid wall region. The axial variations of the enthalpy flux, kinetic energy flux, directed kinetic-energy flux, and momentum flux, all normalized to the mass flow rate of the propellant, are used to investigate the energy conversion process inside the thruster nozzle. The modeling results show that the values of the arc voltage, the gas axial-velocity at the thruster exit, and the specific impulse of the arcjet thruster all increase with increasing hydrogen content in the propellant, but the gas temperature at the nitrogen thruster exit is significantly higher than that for other three propellants. The flow, heat transfer, and energy conversion processes taking place in the thruster nozzle have some common features for all the four propellants. The propellant is heated mainly in the nearcathode and constrictor region, accompanied with a rapid increase of the enthalpy flux, and after achieving its maximum value, the enthalpy flux decreases appreciably due to the conversion of gas internal energy into its kinetic energy in the divergent segment of the thruster nozzle. The kinetic energy flux, directed kinetic energy flux and momentum flux also increase at first due to the arc heating and the thermodynamic expansion, assume their 123Plasma Chem Plasma Process (2010) 30:707-731 DOI 10.1007/s11090-010-9257-0 maximum inside the nozzle and then decrease gradually as the propellant flows toward the thruster exit. It is found that a large energy loss (31-52%) occurs in the thruster nozzle due to the heat transfer to the nozzle wall and too long nozzle is not necessary. Modeling results for the NASA 1-kW class arcjet thruster with hydrogen or decomposed hydrazine as the propellant are found to compare favorably with available experimental data.

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Section: Resultsmentioning

confidence: 99%

Section: Discussionmentioning

confidence: 96%

Section: Modeling Approachmentioning

confidence: 99%

Section: Introductionmentioning

confidence: 99%

See 2 more Smart Citations

Modeling Study on the Flow, Heat Transfer and Energy Conversion Characteristics of Low-Power Arc-Heated Hydrogen/Nitrogen Thrusters

Wang

Geng

Chen

et al. 2010

Plasma Chem Plasma Process

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“…At a flow rate of 41 sccm (4.03 mg.s −1 ), F total is 1.4 mN (G = 1) and remains constant thereafter. In a typical DC arcjet [20] or cascaded arc [21], the electron density is very high (∼10 15 cm −3 ) with the discharge near Local Thermal Equilibrium (neutral gas temperature approaching electron temperature) and the energy transfer is dominated by elastic electron-neutral collisions. In such high electron density mode typical of an arc or a filament there are often issues with stability of the arc and electrode erosion but interesting thruster characteristics have been reported for non-reactive gases such as helium [22].…”

Section: Effect Of Gas Flowmentioning

confidence: 99%

Direct Measurement of Axial Momentum Imparted by an Electrothermal Radiofrequency Plasma Micro-Thruster

et al. 2016

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Gas flow heating using radio frequency plasmas offers the possibility of depositing power in the center of the flow rather than on the outside, as is the case with electro-thermal systems where thermal wall losses lower efficiency. Improved systems for space propulsion are one possible application and we have tested a prototype micro-thruster on a thrust balance in vacuum. For these initial tests, a fixed component radio frequency matching network weighing 90 g was closely attached to the thruster in vacuum with the frequency agile radio frequency generator power being delivered via a 50 ohm cable. Without accounting for system losses (estimated at around 50%), for a few 10 s of Watts from the radio frequency generator the specific impulse was tripled to ∼48 s and the thrust tripled from 0.8 to 2.4 milli-Newtons.

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