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A cylindrical rotating detonation engine (24 mm diameter) with multiple injector holes on the combustor side wall for both propellant injection and cooling by injector gas flow was designed and te...

Low fuel regression rate is one of the major drawbacks of hybrid rocket motors. In this study, a novel injector concept is proposed to provide a substantial enhancement in the fuel regression rate....

It has been urgently required to develop ionic liquid electrospray thruster (ILET) for rapidly expanding micro-/nanosatellites. In this work, ILETs with high-density array emitters are fabricated t...

This study systematically investigated the influence of the position and inclination angle of an externally mounted hollow cathode in 1.35 kW magnetically shielded Hall thrusters to establish the r...

The performance of a three-channel, 100-kW nested Hall thruster was evaluated on xenon propellant for total powers up to 102 kW. The thruster demonstrated stable operation in all seven available ch...

This study provides an experimental analysis of the throttling methods implemented using a throttleable injector that generates a thrust in the range 24–64% on a 50 N monopropellant thruster using ...

Two-fluid atomization of highly viscous and shear thinning non-Newtonian fluids must overcome the viscous resistance. Conventional air-blast atomizers address this requirement by operating at highe...

The plasma plume properties of a three-channel 100-kW-class nested Hall thruster were measured on xenon propellant for total powers up to 80 kW. The thruster was throttled through all seven availab...

Combined with a common fuel binder, solid hypergols can simplify the overall complexity of hybrid rocket engines, as the fuel grain can be ignited and reignited without any external power source or...

Small, low-thrust, long-burn-time solid propellant rocket motors could provide propulsion for a new class of kilogram-scale, transonic, uncrewed aerial vehicles (UAVs). This paper investigates technological challenges of small, low-thrust solid rocket motors: slow-burn solid propellants, motors that have low thrust relative to their size (and thus have low chamber pressure), thermal protection for the motor case, and small nozzles that can withstand long burn times. Slow-burn propellants were developed using ammonium perchlorate and 0–20% oxamide (burn-rate suppressant), with burn rates of 1–4 mm⋅s−1 at 1 MPa. Using these propellants, a low-thrust motor successfully operated at a thrust/burn area ratio 10 times less than that of typical solid rocket motors. This kilogram-scale motor can provide 5–10 N of thrust for 1–3 min. An ablative thermal protection liner was tested in these firings, and a new ceramic-insulated nozzle was demonstrated. This paper shows that small, low-thrust solid motors are feasible and presents a baseline design for the integration of such a motor into a small UAV., Department of Defense (DoD), MIT Lincoln Laboratories, BAE Systems, Inc.

A rotating detonation engine (RDE) generates a continuous thrust with one or more rotating detonation waves. Because of the velocity on the order of kilometers/second, the reaction zone is relatively small. Therefore, the RDE realizes a short combustion chamber length. However, the detonation waves induce an azimuthal motion of propellant, resulting in torque around the thrust axis. Because the motion does not contribute to the thrust, the torque is important in terms of performance loss. Herein, we conducted combustion tests with a six-axis force sensor to simultaneously measure 0.149±0.009 Nm torque and 48.1±0.9 N thrust. A comparison of detonation waves captured by high-speed camera revealed that the torque followed the direction and was offset when the waves existed in both of two directions simultaneously, which indicates the possibility of controlling the torque. Under a mass flow rate at 87±9 g/s and an equivalence ratio at 1.43±0.28, when the azimuthal component of shear force was 8.8±0.6% of the thrust, 0.77±0.10% of the total kinetic energy of the exit flow was distributed to the azimuthal component of velocity and did not contribute to the thrust. We therefore concluded that the effect of the azimuthal motion on the RDE’s performance was small.

This paper describes the testing and pulse performance analysis of a 45 N thruster, designated A45, using a 19.78% monomethylhydrazine–80.22% hydrazine fuel blend with mixed oxides of nitrogen cont...

We present a straightforward formulation predicting the characteristic velocity and specific impulse for bipropellant thrusters as a direct function of injection conditions, propellant combination,...

A reversed gas feed configuration for Hall thrusters is proposed and studied numerically. As an alternative to standard direct injection, we investigate the effect of injecting the propellant near the channel exit and directing it backwards, towards the anode. The resulting neutral density and average velocity fields are studied with the Direct Simulation Monte Carlo method for both cold and warm anode conditions. The residence time of neutral particles inside the channel and in the ionization region is computed using a test-particle Monte Carlo method. The computations indicate that the reversed injection allows for increasing the mass utilization efficiency of a standard feed configuration from 2-5% to a maximum of 20-30% depending on the initial efficiency.

An electrospray thruster designed for nanosatellites is introduced in this paper. The thruster has a compact size featuring a passive propellant transport method through the use of porous materials...

形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2021-07-18, 資料番号: PA2210010000

To further characterize the roles of aluminum and ammonium perchlorate (AP) characteristics in ammonium perchlorate/R45-M hydroxyl-terminated polybutadiene (AP/HTPB)–composite propellants, aluminiz...

Experimental testing of a number of novel additively manufactured monopropellant microthrusters was conducted under atmospheric conditions using 87.5% concentration hydrogen peroxide. The aim of th...

17 USC 105 interim-entered record; under review. The article of record as published may be found at http://dx.doi.org/10.2514/B38282 Typically, the burning surface of a composite solid propellant is controlled through grain geometry and formulation. However, combustion studies of grains constructed from different propellant formulations at fine scales (nominally 1 mm) are not readily accessible in open literature. With additive manufacturing, such configurations can be investigated easily. Propellants with a faster burning inner layer (enhanced with either 1 wt.% iron oxide or 5 wt.% nanoaluminum) were 3D printed between two layers of slower burning 85 wt.% ammonium perchlorate/hydroxyl-terminated polybutadiene propellant. The dynamic combustion behavior of the layered propellant was investigated at pressures ranging from 3.45 to 10.34 MPa. Overall, an increase in the burning surface area, without interlayer delamination, was observed. The driving force behind the propellant surface area increase was the difference in the burning rate between the layers. In addition, the nanoaluminum propellant layer had a more stable burning rate exponent than the cast nanoaluminum propellant. Overall, only a small addition of catalyzed propellant was needed to increase the burning rate of the bulk material. The results of this study lay the foundation for functionally grading propellant grains, which could tailor the thrust profile of solid rocket motors and gun propellants. NASA Space Technology Research Fellowship 80NSSC17K0176

The overall propulsion efficiency for ion-mode electrospray thrusters has been predicted to be as high as 90%; however, experimental measurements currently fall far short of these predictions. Furt...

This study directly measures the performance of a thruster nozzle using water vapor as propellant in the Reynolds number below 1000. Because of its properties, water is suitable as a propellant for...

Conventional expressions and definitions describing performance of plasma thrusters, including the thrust, specific impulse, and the thruster efficiency, assume a steady-state plasma flow with a co...

A variety of methods have been developed to enhance solid propellant burning rates, including adjusting oxidizer particle size, changing the formulation, modifying the propellant core geometry, and...

The Photonic Laser Thruster (PLT) efficiently uses photons as propellant and thus promises to enable innovative space endeavors beyond the reach of propellant-based conventional thrusters. Previous...

Aluminum (Al)/lithium (Li)-alloy-based fuels can potentially improve composite propellant performance and reduce hydrochloric acid formation. Shattering microexplosions have been observed in Al–Li-...

The response behavior of capillary tube attenuated pressure sensor arrangements is studied using a rotating detonation engine as a nonlinear pressure wave source. The resulting measured pressures a...

In this study, to reduce oxidizer to fuel ratio (O/F) shifts, which are a serious problem for a hybrid rocket used for launching a satellite, an altering-intensity swirling-flow-type hybrid rocket ...

Nondeterministic analyses were performed on a one-dimensional, steady, chemically reacting, reduced-order model for hydrogen peroxide flow through a monopropellant catalyst bed. Polynomial chaos ex...

Paraffin-based fuels incorporating solid amine boranes are investigated to identify formulations suitable for use as hypergolic hybrid rocket propellants. Their ignition delays are measured followi...

Room-temperature ionic liquids offer a potential technological approach for propellant production on Mars using in situ resources. These unique salts can be employed to preferentially capture carbo...

Aging induced changes in the mechanical properties of solid propellant can lead to defects, such as cracks and grain–liner separations, which limit the service lifetime of solid rocket motors. This...

A pulsed electric thruster named Pulsed Electrogasdynamic (PEG) thruster was investigated. In this thruster, the propellant gas is accelerated gasdynamically to high exhaust velocities (10–15 km/s...

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

The performance of many Hall thrusters has been shown to exhibit sensitivity to background pressure in ground-based test facilities, and this property can have an impact on the prediction of in-spa...

In recent years, phosphor thermography has received increased attention as an alternative to other common temperature measurement techniques (e.g., thermocouples, pyrometry). This technique has bee...

This paper describes a metal plasma thruster (MPT) that is suited to nano- and microsatellites. The MPT uses solid metal propellant (hence requires no liquids, gases, flow valves, or flow controls)...

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...

This paper focuses on investigating “combustion instability,” a phenomenon that mainly involves the interaction of the propellant flames with the acoustic oscillations prevalent in full-scale rocke...

Subnanosecond laser ablation of asteroid simulants has been studied, and direct force measurements were made. A 0.7 ns pulsed laser was used, with a 40 kHz pulse repetition frequency, 1064 nm wavel...

Iodine has the potential to be a mission-enabling propellant for electric propulsion; however, iodine as the working gas in a hollow cathode presents many challenges. Iodine propellant was compared...

Staging of space propulsion systems would allow lifetime limitations inherent to small propulsion systems to be bypassed in order to enable high-ΔV capabilities for small spacecraft, in particular mass and volume constrained CubeSats. In addition, staging can be used to provide redundancy in the propulsion system, counteract thruster degradation, or open up new avenues of mission optimization. Analytical approximations are developed in order to provide a computationally simple approach to the design, analysis, definition of propulsion technology requirements, and online autonomous decision making for systems that make use of staging of propulsion elements. In addition, the analytical approximations provide insight into the dependencies of performance metrics on system parameters. Equations are developed for any mission, defined by its ΔV, and then specialized for escape trajectories., NASA (Grants 80NSSC18M0045, 80NSSC18K1186)

The multifluid computational framework Hall2De is updated to perform simulations of Hall thrusters operating with krypton propellant. The implementation is validated by comparison to past experimen...

A combustor was developed to operate with natural gas and air as the primary propellants at elevated chamber pressures and air preheat temperatures representative of land-based power generation gas...