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In this paper, the authors investigate whether airbreathing engines have useful application to vertical takeoff and vertical landing systems, which currently represent the mainstream for reusable l...

This paper presents the design methodology and performance testing of an additively manufactured resistojet operating on hydrogen as propellant. Additive manufacturing allows to produce complex monolithic resistors, resulting in reliable high efficiency thrusters. The concept, to be used in combination with advanced cryogenic storage technologies, is proposed for short time and high specific impulse orbit transfers. The simplified two-dimensional thermal design approach adopted is discussed, and its application to the engineering of the resistor is shown for both Inconel 718 and tungsten. The paper reports the performance testing of the proof-of-concept version of the thruster, manufactured in Inconel 718. Experiments on hydrogen show near ideal performance, demonstrating a peak thermal efficiency of 96%. The thruster proves the validity of the design methodology proposed, and the feasibility of the approach to develop monolithic additively manufactured hydrogen resistojets as main propulsion units.

A hollow cathode is an efficient electron source in the self-heating mode utilizing the discharge power. However, in sub-ampere currents, it needs keeper power to maintain the thermionic electron discharge, which could decrease the thrust efficiency. To address this problem, we propose using a microwave cathode, which is based on the flight model of a microwave ion thruster neutralizer cathode, as an alternative to a hollow cathode. First, we redesigned the magnetic field of a microwave cathode discharge chamber and tested it in the diode mode configuration. The electron emission current is doubled compared to the original performance. Next, we coupled the improved microwave cathode with a 200-W class Hall thruster and compared the characteristics and performance with a hollow cathode. We confirmed that the magnetic field polarity affects the ignition characteristics. We measured the thrust by an inverted thrust stand, the ion energy distribution functions by a retarding potential analyzer, and the beam profiles by an ion collector. The thrust and thrust efficiency are equivalent for both types of cathode. The specific impulse is 10% higher in the case of the microwave cathode. Since the potential difference between the microwave cathode and ground rapidly increased at currents above 600 mA, this could be taken to be the trade-off point against the hollow cathode., [Highlights] / • Electron current of a microwave cathode doubled by redesign of magnetic circuit. / • Magnetic polarity of microwave cathode and Hall thruster affected plasma maintenance. / • Thrust was equivalent in the both cases of microwave cathode and hollow cathode. / • Performance trade-off point between the two cathodes was about 600mA. / • Efficiency analysis by IEDF and beam profile showed different plasma characteristics.

形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2020-08-25, 資料番号: PA2010026000

A low-cost and reliable propellant flow controller is necessary to offer competitive electric propulsion systems. This paper first reviews xenon flow control methods for Hall thrusters and ion engi...

Accepted: 2019-09-18, 資料番号: SA1190110000

In this study, a novel single-piece thin multi-layer tungsten resistive heater was successfully fabricated using additive manufacturing and tested as an electrothermal thruster. The heater has 12 resistive layers, with each layer having a thickness and height of 0.15 and 81 mm, respectively, and can provide high heating efficiency. A single-piece or monolithic heater was manufactured via additive manufacturing technique, which drastically improved its reliability and decreased its manufacturing cost. In the heating and thrust measurement tests that used nitrogen gas as a propellant, the heater reached a gas temperature of ∼2000 K at a 140-A heater current without experiencing any failure. The tungsten-heater resistance linearly increased with an increase in temperature due to the temperature dependence of tungsten’s resistivity. The specific impulse and thrust increased with the heater temperature in accordance with the theoretical prediction. Even including a voltage drop due to a contact resistance, the achieved heater efficiency reached 63% at a 100-A heater current even without a thermal insulation around the thruster. The heater efficiency decreased with an increase in the heater temperature due to heat loss to the surroundings. The heat-loss analysis indicated that both thermal conduction and radiation heat losses were crucial for improving the heater performance at a high-temperature operation of over 2000 K.

Accepted: 2018-05-21, 資料番号: SA1180213000

形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2018-09-19, 資料番号: PA1910034000

An advanced background-oriented schlieren (BOS) method, named as the speckle beam-oriented schlieren technique, was newly developed to measure the distribution of refraction angles in transparent media. A speckle pattern is generated by passing a coherent laser beam through a holographic diffuser, a pinhole, and a lens, generating a collimated background image that is projected directly onto the image sensors. Since the intensity of the background image is maintained at a high level, this method is, in principle, useful for diagnosing fast and/or low signal-to-noise phenomena, such as high-temperature gasses with radiation emission. Moreover, by splitting the background beam into two imaging paths with different focal lengths, the refraction angles can be measured for a schlieren object with uncertain location, and the depth position of the refraction angles can be resolved. This technique was demonstrated by measuring the refraction angle and the depth position distribution in a sonic jet with different injected locations., Published: 08 January 2021 ファイル公開：2022/01/08

Nanosecond-pulsed surface dielectric barrier discharge plasma actuators are used to control shock-wave/boundary-layer interactions. Characterization of the separation region without plasma actuatio...

This paper reports measurements of the xenon ground state and excited state densities inside a μ10 microwave ion thruster. This thruster exhibits a 40% thrust enhancement upon changing from the waveguide to the discharge chamber propellant injection mode. In the present work, the associated mechanism was quantitatively evaluated using two-photon laser induced fluorescence (TALIF) spectroscopy to monitor the thruster waveguide. The 834.7 nm emission from excited state xenon was investigated with a 224.3 nm dye laser to excite the Xe I 5p61 S0 6pʹ [3/2]2 state, compared with the emission without the laser. The resulting data confirm that the neutral density exhibits a linear relationship with the propellant flow rate in the cold gas and ionized state, while the ion acceleration decreases the neutral density by the same order of magnitude as the propellant utilization efficiency is changed. As the propellant flow rate increases, the collisions of neutrals that generate excited states occur in the waveguide and, when this process plateaus, the ground state emission suddenly increases. Propellant injection from the discharge chamber is evidently effective at suppressing collisions with electrons in the waveguide that generate excited states and that potentially interfere with microwave propagation.

Accepted: 2019-07-28, 資料番号: SA1190155000

The thermal behavior of cryogenic propellant tanks is crucial issue in the operation of cryogenic propulsion systems. Herein, ground experiments were conducted in a 600-mm-diameter cryogenic tank filled with liquid nitrogen. As pre-pressurants, gaseous helium and gaseous nitrogen (of the same species as the liquid), were used to investigate its effect in accordance with actual propulsion systems. The tank was sealed after pre-pressurization to observe the self-pressurization. The evaporation rate and heat flow in the tank were estimated based on pressure and temperature measurements. In addition, the axial liquid temperature distribution was obtained through the liquid draining from the tank bottom, and a thermal stratification model was developed. These results demonstrated that the type of pre-pressurant significantly affected the thermal behavior in the tank. A higher evaporation rate and higher liquid internal energy rise rate with a thicker thermal layer were observed in the helium pre-pressurization case. The lower nitrogen partial pressure in the helium case enhanced the vaporization and growth of the thermal layer. Estimation of the power balance in the tank demonstrated that not only the ullage but also the heat mass of the tank provided heat for the evaporation and thermal layer. The evaporation occurs mainly at the contact point between the tank skin and liquid surface. The nitrogen vapor rises in a thin layer along the tank skin because of buoyancy under nitrogen pre-pressurization; however, buoyancy is lower under helium pre-pressurization, and a radial vapor flow is probably produced from the contact point instead, leading to a higher heat flux to the liquid.

Accepted: 2019-11-06, 資料番号: SA1190188000

A 3D-printed resistojet, wherein the heater and flow path were united, was produced. The heater consists of multi-layer shells and is difficult to break down. In this paper, the thermal design and the results of the thrust measurement are reported. The measured electrical resistance matched with the predicted one. Both the heater and the thrust reached 70 percent efficiency when paired with a nitrogen propellant at a mass flow rate of 0.2 g/s., AIAA 2018-4907 Session: Micropropulsion II

Accepted: 2018-05-25, 資料番号: SA1180088000

Accepted: 2014-07-21, 資料番号: SA1150044000

Accepted: 2012-02-14, 資料番号: SA1004276000

This study focuses on understanding and modeling the physical phenomena that occur in degraded zones of silica-phenolic (SiFRP) materials under exposure to high-temperature gasses when applied to a liquid rocket engine (LRE) combustor. Although understanding and modeling these phenomena is considered essential in designing an LRE combustor, few studies on these fields can be found in the available literature. Basically, it is well known that when ablators are heated, a pyrolysis reaction proceeds in them, forming three distinct zones: a charred, a decomposed, and a virgin zone. The obtainable information for the thermal response of SiFRP in ground-firing tests is classified in two categories. The first category involves the equilibrium state characteristics after a long time has elapsed following burnout. This refers to the degraded thickness distribution, which reflects 3D information (the combustor’s inner surface x the thickness direction) regarding the heat load distribution over the entire combustor’...

資料番号: PA1210040000

Ultimate weight reduction and low manufacturing cost are strongly required in the development of next-generation launch vehicles. It is one of effective methods for these purposes to change structural materials from conventional aluminum alloys to composites, such as CFRP (Carbon Fiber Reinforced Plastics). Under these situations, we developed a launch vehicle structure by using CFRP/honeycomb sandwich panels in order to achieve both weight and cost reduction. FEM (Finite Element Method) simulation was utilized to effectively carry out the structural design. Metallic structures are applied for the flanges (joining interfaces between the other structures) in general; however, we used CFRP structures also for the interface flanges for the weight reduction and manufacturing simplicity. The designed CFRP structure was fabricated and a static load test was conducted considering flight loads during actual launch. The strain distribution was obtained by FBG (Fiber Bragg Grating) sensors and three-dimensional displacement was measured by laser tracker as well as normal strain gauges in the limit load test. These measured data were compared with the FEM simulation results and they show good agreement. We will apply these manufacturing, testing, simulation and sensing techniques for the future launch vehicle developments.

This study numerically explores the flow physics associated with nanosecond pulsed plasma actuators to control shock-wave induced boundary-layer separation. By using two actuators, parallel and canted with respect to the main flow direction, a previous experiment suggested the actuator worked in two ways: boundary layer heating and vorticity production. The heating effect was enhanced with the parallel electrode and made the separation stronger, while the canted electrode produced vorticity and suppressed the separation due to the momentum transfer. Because the detailed physics is still unclear, a numerical investigation is undertaken with a large eddy simulation and an energy deposition model for the actuator. The flow without the actuation corresponds to the experimental observation, indicating the calculation successfully resolves the separation. With the actuation, as with the experiment, the calculation successfully demonstrates definite difference between the parallel and canted electrodes: the parallel electrode causes excess heating and increases the separation, while the canted electrode leads to a reduction of the separation, with a corresponding thinning of the boundary layer due to the momentum transfer. The counter flow created by the canted actuator plays an important role in the vortex generation, transferring momentum to the boundary layer and, consequently, mitigating the separation., Accepted manuscript for publication in Physics of Fluids