Showing papers on "Arcjet rocket published in 2016"

Chemical Nonequilibrium Modeling of Low-Power Nitrogen/Hydrogen Arcjet Thrusters

Abstract: Departures from thermal, chemical equilibrium of a plasma flow in a low-power arcjet thruster with a 1:2 nitrogen/hydrogen mixture as the working gas have been studied by axisymmetric numerical simulations. Electrons, ions, atoms, and molecules are represented as separate chemical species in the plasma mixtures of nitrogen and hydrogen. The predicted temperatures and densities of the electrons, heavy particles, and plasma velocity are presented throughout the whole flowfield of the nozzle. Significant temperature discrepancies between electrons and heavy species have been found in the near-anode wall region, which indicates that thermal nonequilibrium largely controls the near-anode electron densities and furtherly governs current attachment on the anode. The distributions of nitrogen and hydrogen species inside the arcjet thruster are also examined. It is found that the enrichments of hydrogen species at the center of the thruster and anode surface are due to different reasons. Some important kinetic pro...

Radiative Transfer In A Rigid Carbon Material Under Arcjet Flow Condition

Abstract: Radiative transfer within a rigid carbon material in a nitrogen arcjet heating condition is studied experimentally and numerically. The experiment is done using a 1MW constrictor-type arcjet wind tunnel. Nitrogen is used as a test gas to minimize the surface recession of the carbon material. The nitrogen arcjet freestream is calculated from the upstream end of the constrictor up to the test section of the wind tunnel using an integrated numerical method developed recently. The thermal response analysis of the heated carbon material is made by a coupled manner between the arcjet flows and the material. The radiative conductivity is used in the radiative transport calculation. The radiative extinction coefficient to obtain the radiative conductivity is evaluated by using X-ray CT models. Comparison of the surface and in-depth temperature between calculation and measurement is made. The result shows that a reasonable agreement is obtained by accounting for the radiative transfer.

Design features and experimental researches of an arcjet thruster for small satellite

Abstract: The direction of solution of actual problems of further improvement of the small satellites with propulsion system (PS) by designing arcjet thruster with low power consumption (up to 70 W) and high specific impulse (up to 350 s) was considered on the basis of the research on the maneuvering smallsat with ammonia corrective propulsion system with Electroheat Microthrusters (EHMT). The purpose of the research is confirming technical possibility of designing an arcjet thruster for smallsat with power consumption 60…70 W and specific impulse 300…350 s. In the course of research tasks on determination of construction appearance of a demo sample of No. 2 of arcjet thruster, exploration of its temperature characteristics defining gauge of specific traction impulse of a study of erosion resistance of thruster cathode material at impact of an electric arc are resolved. A constructional scheme of arcjet thruster with visualization of an electric arc is developed, demo arcjet thruster samples are made and experimental arcjet thruster research is conducted in operation on nitrogen, argon. Working fluid temperature in arcjet thruster nozzle behind critical section is reached 1400 K at 68 W, which allows to predict specific impulse value of arcjet thruster in operation on ammonia in 300…350 s. On basis of experimental researches of erosion resistance of arcjet thruster cathode material (tungsten) at impact of an electric arc the necessity of the further research is shown for substantiation of a technical possibility of design arcjet thruster in composition of PS smallsat with required resource of operation. The experimental researches is the continuation of the works in the field of design of EHMT of various construction and represents practical interest during creation of PS for maneuvering smallsat.

Analysis of Air Plasma Flows in Magnetoplasmadynamic Arcjet Testing

Abstract: The influence of the electrical input power of a magnetoplasmadynamic arcjet on the flow translational temperature and velocity is investigated. From a nominal testing condition, representing the 78.8 km altitude trajectory point on the Hayabusa re-entry, the electrical power is varied from 146 kW to 189 kW and the temperature and velocity are measured using a system coupling a spectrometer and a Fabry-Perot interferometer to determine the Doppler broadening and Doppler shift of the atomic emission lines.The atomic nitrogen translational temperature is measured using four different emission lines resulting in a median temperature value of 10700 K for the nominal test condition. The atomic oxygen temperature is measured using the 777 nm triplet with a median temperature value of 16350 K. A reason for the higher temperature of the oxygen has not been determined, however, this is in line with results previously obtained on a different plasma generator.The temperatures of both the nitrogen and oxygen increase approximately linearly with electrical input power. The effect of the power increase is significantly larger on the nitrogen temperature with a measured increase of 38 K / kW (0.35 % / kW) compared to the oxygen increase of 24 K / kW (0.15 % / kW). This is thought to be due to the direct energy transfer from the arc into the nitrogen whereas the oxygen is energised through particle collisions downstream of the arc.The measured velocities were the same for nitrogen and oxygen. For the nominal condition the velocity was measured at 3350 m / s. The velocity also increases approximately linearly with the electrical power increase. The velocity increase was determined to be 22 m / s / kW (0.66 % / kW). When considered as a percentage this is a stronger effect than seen in the temperature increases.

Theoretical and Experimental Investigations on rocket Propulsion of Counterprojectile of Active Protection System

Abstract: The paper presents selected results of a research project the aim of which is to perform the technology demonstrator of an active protection system. One of the elements of this system is a smart counterprojectile. The counterprojectile head includes electronic components sensitive to high overload. To avoid such overload effects, a solid propellant rocket engine was used as a propulsion system. The design concept of a propulsion system of a counterprojectile is described. Due to short operation time, it is assumed that the propellant system will have the features of a rocket booster. A propellant charge will consist of seven rocket propellant grains with low-combustible layer thickness and a large area of burning, ensuring high-power engine thrust. Based on the Polish new homogeneous solid rocket propellant, and adopted dimensions of the propellant charge and nozzle block, calculations of ballistic characteristics of the propulsion system were carried out. In order to verify the results of a theoretical analysis, experimental studies were carried out.

Arcjet propulsion systems for spacecraft

Abstract: An arcjet thruster system for a spacecraft is provided. The arcjet thruster system may include a power supply that includes a radio-frequency start power supply and a continuous direct-current power supply, each selectively coupled to electrodes of an arcjet for initiation and maintenance of an arc between the electrodes. A radio-frequency/direct-current control module may be provided for selectively coupling the radio-frequency start power supply and a continuous direct-current power supply. The radio-frequency start power supply may be used to initiate an arc that is then sustained by the continuous direct-current power supply.

우주비행체 추진기관 기술 현황 및 전망

Abstract: Spacecraft propulsion system is a kind of rocket engine that has been developed from the end of 1950s for attitude control and orbit maintenance of satellite. Since the spacecraft propulsion system has to be used for a relatively long time, therefore, stability of propellant and life of thruster could be very important factor for propulsion system design. Recently, green propellant propulsion and all electrical propulsion system have became very important issue, and we also need a development according to well organized plan. In this paper, we will introduce the development status, key technologies and development prospect of spacecraft propulsion system.

Velocity Measurements in an Arcjet Erosion Test Facility

Abstract: The arcjet test facilities at Arnold Engineering Development Complex include the capability to perform particle erosion testing in high-enthalpy flow. Carbon particles are injected into the arc heater upstream of the Mach 3.5 exit nozzle. The particles are entrained in the flow and accelerated to supersonic speeds. Calculations have predicted that the carbon particles do not reach freestream flow velocities due to the relatively short acceleration path in the arcjet facility. Two nonintrusive diagnostic techniques were used to measure the particle speeds. This paper describes the test apparatus and presents the calculated and measured freestream and injected particle velocities. Measurement results indicate that the particle velocities are approximately 1,900 to 2,100 m/s and freestream air velocity is approximately 3,100 m/s.

Response of the Arc Plasma Source to Combustion Chamber Pressure Fluctuation of the Small Size Thruster Using Plasma Support Combustion

Abstract: This paper deals with the influence of variation in combustion chamber pressure on an arcjet plasma source for an arcjet-assisted thruster. Arcjets have been applied to chemical thrusters in order to promote combustion and augment performance. Some groups reported that combustion of solid propellant and monopropellant such as SHP163 was successfully sustained with arcjet. Arc discharge is, however, negatively influenced by variation in combustion chamber pressure. Hence, we propose to apply active control to arcjet-assisted chemical thruster in order to stabilize combustion. The controller design requires the response of arcjets to variation in combustion chamber pressure variation. In this study, we investigated the influence of step-like pressure change on arcjet exit using a combustion chamber simulator. In the combustion chamber simulator, pressure was suddenly increased from 0.1 to 0.35 MPa using a nitrogen-filled buffer with a burst diaphragm. For a pressure rise of 0.15 MPa with a time constant of 0.2 s, arc discharge was interrupted immediately after sudden rise in pressure of the combustion chamber simulator. From the results, arcjet was negatively affected by the combustion chamber pressure.