The transition from the low-amplitude, aperiodic fluctuation to thermoacoustic instability is usually treated as a Hopf bifurcation, but the transition incorporates complex dynamical behaviors. Intermittent oscillations commonly occur in the transition process wherein chaotic states alternate with limit cycle oscillations and both the chaotic state and the limit cycle oscillation last with no fixed

Journal of Propulsion and Power, Ahead of Print.

We conduct a high-speed visualization of coolant liquid film dynamics inside a 10N-class bipropellant thruster using monomethylhydrazine and a mixture of nitrogen tetroxide with approximately 3% nitric oxide as propellants at equivalent conditions with the flight model. The direct visualization of the liquid film inside a quartz glass chamber demonstrates the transient dynamics of film flow from ignition

The response and performance of an aeroelastic hypersonic intake was studied experimentally using fundamental geometry and structural boundary conditions. The experiments were conducted in a hypersonic wind tunnel at a Mach 5.85 condition. The most relevant deforming component was the compression ramp, which was treated as a cantilever surface to emulate the global deformation of the intake. Flowfield

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

Closer integration between the airframe and the propulsion system is expected for future aircraft to reduce fuel consumption, emissions, weight, and drag. The use of embedded or partially embedded propulsion systems will require the use of complex intakes. However, this can result in unsteady flow distortion, which can adversely affect the propulsion system efficiency and stability. Relative to conventional

The electron-cyclotron resonance thruster with magnetic nozzle relies on two successive energy transfer processes: first from electromagnetic energy to electron thermal energy, facilitated by a coupling structure; and second from electron thermal energy to ion directed kinetic energy, facilitated by a diverging magnetic field. The nature and geometry of the coupling structure are crucial to the first

A high-efficiency concentric tubular-type resistojet with potential application to short-term orbit-raising maneuvers has been fabricated by 3-D printing and demonstrated. The propellant flows through multiple layers of cylindrical shells, with this structure also functioning as a single-piece heater. A 6-cm-high cylinder was realized with a wall thickness of 0.2 mm, using Inconel 718. A nodal thermal

Hybrid–electric powertrains offer the potential for performance improvements in unmanned aerial systems. However, for small unmanned aerial systems, potential gains in range and endurance can depend significantly on the aircraft flight profile and powertrain control logic. Subsequently, these impact the performance of individual powertrain components. This study uses dynamic simulations of an unmanned

Distributed aircraft propulsion has renewed the interest in power-dense, high-efficiency power packs. Ceramic turbomachinery could be a major enabler, although no successful design has been achieved in microturbine rotors. Rotor blade loading is tensile and a hurdle for successful conversion to ceramics. The inside-out ceramic turbine (ICT) rotor uses the superior compressive properties of monolithic

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. A tubular injector, Distributed Tube Injector (DTI), is inserted in the center of the cylindrical fuel port in order to tailor the oxidizer flow introduced into the combustion chamber with the desired combination

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 external fluid. Also, with the hypergolic additive embedded in the binder, the flame zone could be placed at the surface of the grain itself, thereby providing heat to the grain to maximize regression rate

This experimental study investigates low power (3 W) laser ignition and the subsequent burning characteristics of consolidated Al/CuxO nanoparticle pellets and loose nanolaminate flakes. The effects of laser absorption, packing density (bulk porosity), and material properties on ignition are quantified, while simplified numerical models are used to examine their predominance. It is found that ignition

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 to quantitatively study the electrospray emission process of three typical ionic liquids (ILs). Onset voltage, current, voltage–current characteristics, and ratio of positive to negative currents are analyzed

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 available channel combinations for conditions spanning 300 to 500 V discharge voltage and three discharge current densities. A plasma diagnostics array, which included a Wien filter spectrometer, a retarding potential

In this work, two representative hybrid rocket injector designs, a single-port and an axial showerhead design, were examined experimentally to determine the comparative influence on the development of the downstream combustion zone with polymethylmethacrylate (PMMA) fuel. One-dimensional laser absorption tomography was employed to obtain the thermochemical structure of the PMMA–gaseous oxygen reaction

As the level of aerospace technology continues to advance, the requirements for aircraft engines are becoming more onerous. A new thermal-cycle propulsion-technology scheme is proposed based on the existing propulsion system to meet the needs of round-trip space and multitask flights that use hypersonic vehicles, where thermoelectric materials are introduced into the wall of the engine tail nozzle

An experimental study is conducted to advance the understanding of flow physics associated with a boundary-layer ingesting, distributed propulsion system. The influence of incoming boundary-layer thickness on the performance of the system is examined. The propulsion model, integrated with electrical fans, is mounted on a flat plate and tested at subsonic speeds. Detailed characterization of the incoming

Particle image velocimetry measurements have been carried out in a low-pressure turbine cascade operating under unsteady inflow to deeply investigate reduced frequency and flow coefficient effects on flow dynamics, and, consequently, on loss generation in the boundary layer and in the core flow region. Two independent measuring setups have been used for the purpose. The first one captured a large view

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 relationship between thruster performance, plume state, and cathode placement. For this purpose, a combined motor driver with a three-degree-of-freedom movement was designed to move the cathode over a wide

Scramjet engines are one of the most promising hypersonic airbreathing propulsion technologies for robust, efficient, and economical access to space. Multi-objective design optimization has been conducted for Busemann-based intakes in inviscid and viscous regimes at single and multiple design points, respectively, by means of surrogate-assisted evolutionary algorithms coupled with computational fluid

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 channel combinations at discharge voltages from 300 to 500 V and three different current densities. The resulting test matrix contained forty-six unique conditions ranging from 5 to 102 kW total power and 16

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

We present a straightforward formulation predicting the characteristic velocity and specific impulse for bipropellant thrusters as a direct function of injection conditions, propellant combination, and nozzle expansion ratio. The theoretical formulation deduces a framework for a quantitative noncombustion test or cold-flow test to predict the performance indices. We simply employ water and dyed water

The lack of established optimal design guidelines for turbomachinery operating in the nonideal flow regime (e.g., organic Rankine cycle turbines, CO2 compressors, compressors for refrigeration systems) demands for effective and efficient automated design methods. Past research work focused on gradient-free methods applied to computational fluid-dynamic simulations. The application of the adjoint method

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 containing 3% nitric oxide as the oxidizer. The thruster components are fabricated using additive manufacturing (AM), which enabled fine design features on injectors to achieve fast response times. The main focus

This work presents a novel approach for the modeling of the entrainment in the numerical simulation of the internal ballistics of hybrid rocket engines with paraffin-based fuels. This model, coupled with a more sophisticated gas–surface interaction treatment, is an improvement of the model previously developed by some of the authors, which was based on some oversimplifying assumptions. Indeed, the

A subscale, research rocket thrust chamber operating with cryogenic oxygen and hydrogen exhibits self-excited transverse-mode instabilities with amplitudes of more than 80% of the steady combustion chamber pressure (peak-to-peak) for some operating conditions. During unstable combustion, an increase in the integral heat flux into the water-cooled combustion chamber walls of 20–40% with respect to stable

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 this work was to select a specific catalyst bed geometry for the thruster system and to investigate more general methodologies for monopropellant packed catalyst bed optimization. Characteristic velocity efficiencies

Real-time detection and control of the isolator shock-train leading edge (STLE) is important to the performance of high-speed air-breathing engines, such as dual-mode scramjets. Typically, the STLE location is determined using wall static-pressure measurements, but there are often restrictions on the placement and overall number of the pressure transducers, reducing the viability and accuracy of such

Two types of combustion instabilities were observed in a multi-injector research combustor with liquid oxygen and liquefied natural gas (LOX/LNG) propellants operating at conditions similar to an upper stage rocket engine. The first was identified as an excited tangential acoustic mode of the combustion chamber and can be explained by coupling with the LOX injectors, similar to the mechanism previously

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 higher gas–liquid mass ratios (GLRs) at the cost of deteriorated efficiency. In this investigation, the atomization efficiency ηatom of a two-fluid atomizer is characterized. Fluidic design of the atomizer reduces

Low-thrust long-burn-time solid rocket motors may be useful as propulsion for small, fast, uncrewed aerial vehicles. These motors require a slow-burning propellant that can operate at unusually low chamber pressures (0.3–2 MPa). Slow-burn propellants were developed using ammonium perchlorate oxidizer and the burn rate suppressant oxamide. By varying the amount of oxamide (from 0 to 20%), burn rates

Electric aircraft propulsion is a growing research area that looks into achieving propulsion through fully electric or hybrid electric systems while achieving low CO2 emissions. The system-level benefit gained by different electric and hybrid-electric propulsion schemes depends heavily on the performance of system-level components in the electric drive-train, including the electric motor, gear box

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. Porous electrospray emitters with 1, 25, and 100 emission tips were fabricated using conventional machining techniques. The geometrical features of the emission tips were measured using a high-precision

A generalized flat plate cascade model is used in this study to predict the unsteady aerodynamics of a vibrating blade row with nonuniform blade spacing in subsonic compressible flow. The blade row is assumed to vibrate in an isolated family of blade-dominated modes. The effect of nonuniform blade spacing on compressor rotor flutter stability is demonstrated by case studies based on the geometric and

Journal of Propulsion and Power, Ahead of Print.

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. Further complicating this is that for passively fed electrospray thrusters, the mass flow rate, which is required to obtain the propulsion efficiency and specific impulse, is not directly controlled or measured

The present research work introduces a novel combustion model formulation based on the flamelet concept suitable to investigate characteristic flow conditions in rocket engine thrust chambers. The new approach extends the validity domain of existing flamelet models to fully cover the observed flow conditions by incorporating kinetic rate effects into a state-of-the-art nonadiabatic flamelet manifold

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 90 wt.% hydrogen peroxide and MnO2/PbO/Al2O3. The injector proposed in this study is a dual-plate injector. It is capable of controlling the mass flow rate using radial slotted and showerhead-type plates

A computational and experimental study was conducted on nozzle geometries for rocket application rotating detonation engines (RDEs). Three geometries, including a nozzleless blunt body typically employed in RDE combustor hot-fire testing and two aerospike nozzles, were investigated. Simulations of the exhaust flow of a rotating high-frequency, high-pressure ratio wave based on rocket RDE test results

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

The combustion of pyrotechnic delay composition based on metallic fuel and perchlorate oxidizer involves complex reaction mechanisms and condensed phase heat release. Studies on the effect of aging on ignition/burning properties, and understanding of such multicomponent pyrotechnic delay mixtures are quite limited. The present work focuses on investigating the reaction kinetics, ignition, and combustion

To further characterize the roles of aluminum and ammonium perchlorate (AP) characteristics in ammonium perchlorate/R45-M hydroxyl-terminated polybutadiene (AP/HTPB)–composite propellants, aluminized and nonaluminized formulations were tested at very high pressures, up to 68.9 MPa (10,000 psi). A total of nine formulations were examined: seven nonaluminized with varying AP distributions, concentrations

Spark discharges were parametrically examined for a cylindrical air-gap electrode configuration. The aim of this study was to elucidate spark discharge characteristics, arc penetration, and exhaust plume development to guide designers of relevant ignition devices. Spark gaps ranged from 0.5 to 2.3 mm, nominal pressures ranged from 150 to 2200 kPa, and two exciter types (bipolar and unipolar) were tested

Carbonaceous deposits can reduce heat flux in fuel-film-cooled kerosene rocket engines, but the deposition process at rocket conditions is poorly understood. Heat flux was measured in a 4.8 MPa (700 psia) fuel-film-cooled kerosene–hydrogen peroxide axisymmetric rocket combustor using 50 null point calorimeters and the Gauss–Newton algorithm with a constant Jacobian matrix. Carbon deposition caused

Experiments have been carried out to investigate the effect of cowl length variation on performance characteristics of a single expansion ramp nozzle. The performance parameters were estimated for cowl lengths of 0, 25, 50, 75, and 100% with respect to the horizontal length of the ramp. Experiments were conducted for different nozzle pressure ratios ranging between 1.5 and 9. The wall static pressure

In this study, a 1.35 kW HEP-100X Hall thruster is used to explore effective methods for improving the global efficiency of a Hall thruster across a wide range of flow rates, from the perspective of optimizing the channel configuration. Experimental and simulation results indicate that an appropriate increase in the length of the straight segment in the variable cross section channel can increase the

An experimental investigation has been conducted on a subscale dual-bell nozzle to study the effect of throttle settings on sneak-transition side loads. The results are from a cold-gas test campaign conducted in a high-altitude simulation chamber for three throttle settings of 1, 2, and 3 bar/s. An analysis is carried out by correlating the simultaneously acquired side-load measurements and real-time

Fuel-film cooling is necessary to mitigate heat transfer in high-pressure oxygen-rich staged combustion engines. A 4.8 MPa (700 psia) axisymmetric kerosene–H2O2 combustor used fuel-film cooling to deposit carbonaceous material on removable metal samples. Posttest inspection of the samples revealed a two-layer structure, with a tenacious dense lower layer and a soot-like upper layer. Total deposit depth

The impact on range and endurance of increasing electrical loads on aircraft and burgeoning interest in all-electric aircraft are driving a need for more efficient replacements for engine-driven mechanical generators. This Paper presents a model of a synergistic gas turbine/solid oxide fuel cell system for combined propulsion and electrical power on aircraft. It is developed using Numerical Propulsion

Journal of Propulsion and Power, Ahead of Print.

Journal of Propulsion and Power, Ahead of Print.

Journal of Propulsion and Power, Ahead of Print.

The Paper involves experimental and numerical heat transfer analyses of four pin-fin cooling configurations in a rectangular channel. Two design concepts were studied with two separate pin-fin (cylindrical and triangular) geometries. The first design includes a single-wall pin-fin configuration where the coolant flows through an array of full-length cylindrical (PF) and triangular pin fins (TF). The

A torsional thrust stand, calibrated for impulse bits in the range of 0.1–0.5 mN⋅s, was used to measure impulse bits from a metal plasma thruster. Impulse data were obtained on roughly 1400 shots, with metal plasma thruster targets of molybdenum, niobium, palladium, aluminum, and carbon. Model predictions (based on a simple circuit model and published plasma parameters) were validated by data from

Lifetime assessment of hollow cathode (HC) sources is an experimentally demanding process, due to the large amount of hours space qualification requires. In this Paper, we present a one-dimensional orifice model coupled to a global-averaged insert model, developed for lifetime prediction of a HC neutralizer in terms of orifice erosion; a modified sputtering yield curve is derived, similar to a previous

This Paper discusses the results of an experimental study to investigate the performance of a converging nozzle with fluidic thrust vectoring, using secondary-jet injection, through the measurements of various parameters such as thrust, flow rates, and pressures over a range of operating conditions. The thrust coefficient is found to be dependent on the nozzle pressure ratio (NPR) of the primary nozzle

A group of liquid simulants was developed in order to ease the process of cold-flow testing by lowering associated costs and hazards while providing accurate analysis of desired liquid behavior through an injector. Simulants were made with the intention of modeling certain physical properties of the liquid such as density, viscosity, and surface tension. These simulant solutions are capable of modeling

This study investigated the impact of fuel injection distribution on flame stabilization and heat release in a cavity-stabilized scramjet. Experiments were conducted using a hydrogen air heater that provided vitiated air at an average stagnation temperature of 1570 K and stagnation pressure of 12.7 atm, accelerated to a Mach number of 2.2 at the combustor entrance. Ethylene was injected using two fuel