Exit Plane Research Articles

The Role of the Centerbody Wake on the Precessing Vortex Core Dynamics of a Swirl Nozzle

Abstract The precessing vortex core (PVC) phenomenon in swirling jets is a helical instability in the flow driven by the coherent precession of the vortex breakdown bubble (VBB) around the flow axis, resulting in the helical rollup of the shear layer. This instability is driven by flow processes in the region upstream of the VBB. Centerbodies, commonly employed in combustor nozzles, create a centerbody wake recirculation zone (CWRZ) that can interfere with VBB precession and hence suppress the PVC. We study this phenomenon in a swirl nozzle with a centerbody whose end face is flush with the nozzle exit plane, using large eddy simulations (LES) and linear hydrodynamic stability analysis for flow Reynolds numbers Re = 48,767 and 82,751, based on nozzle exit diameter and bulk flow velocity. For one of the Re = 82,751 cases, the centerbody end face diameter is halved, resulting in the onset of coherent VBB precession. Linear stability analysis reveals a marginally unstable mode in this case. The same mode is found to be stable in the nominal cases. Structural sensitivity analysis shows that the VBB precession eigenmode is sensitive to changes in the time-averaged flow in the VBB-CWRZ merger region. This suggests that the reduction in CWRZ length due to halving the centerbody end face diameter is the reason for the onset of VBB precession. These results suggest that in general, spatial separation between the CWRZ and VBB can result in the onset of VBB precession and the emergence of PVC oscillations in flows with swirl.

Low power thrust measurements of the water electrolysis Hall effect thruster

We propose that a Hall effect thruster could be modified to operate on the products of water electrolysis. Such a thruster would exploit the low cost and high storability of water while producing gaseous hydrogen and oxygen in-situ as they are required. By supplying the anode with oxygen and the cathode with hydrogen, the poisoning of the cathode is mitigated. The water electrolysis Hall effect thruster (WET-HET) has been designed to demonstrate this concept. The dimensions of the WET-HET have been optimized for oxygen operation using PlasmaSim, a zero-dimensional particle in cell code. We present the first direct thrust measurements of the WET-HET. A hanging pendulum style thrust balance is used to measure the thrust of the WET-HET while operating in the Boltzmann vacuum facility within the Imperial Plasma Propulsion Laboratory. For this test the beam was neutralized using a filament plasma bridge neutralizer operating on krypton. We find thrust, specific impulse, and thrust efficiency all increase linearly with power for values between 400 and 1050 W. Increasing the mass flow rate from 0.96 to 1.85 mg/s increases thrust at the expense of specific impulse. Changing mass flow rate was found to have little impact on the thrust efficiency over this range. An optimal radial magnetic flux density of 403 G at the exit plane is found. Further increases to the magnetic field beyond this point were found to decrease the thrust, specific impulse and thrust efficiency, whereas the discharge voltage increased monotonically with increasing magnetic field for a given input power. It was found that the experimental thruster performance was lower than the simulation results from PlasmaSim. However, the general trends in performance as a function of power and propellant mass flow rate were preserved. We attribute a portion of this discrepancy to the inability of the simulation to model the energy absorbed by the covalent bond of the oxygen molecule. For the powers and mass flow rates surveyed we measured thrust ranging from 4.52pm 0.18, to 8.45pm 0.18,mN, specific impulse between 324pm 12, and 593pm 12,s, and anode thrust efficiencies between 1.34pm 0.10, and 2.34pm 0.10,%.

High-spatial-resolution image reconstruction-based method for measuring electron temperature and density of the very near field of an applied-field magnetoplasmadynamic thruster

A method based on image reconstruction is proposed for measuring the electron temperature () and density () distribution of the axisymmetric low-temperature argon plasma plume of applied-field magnetoplasmadynamic thrusters (AF-MPDTs). The method uses a complementary metal–oxide–semiconductor camera with narrowband filters of 460 and 500 nm to obtain the raw projection images in specific wavelength ranges. An image reconstruction method is developed to construct relative emission intensity profiles. The electron temperature distribution can be obtained by combining the relative emission intensity ratio of two profiles with a simple argon kinetic model. The electron density distribution can be calculated by constructing a function of relative emission intensity containing and . To verify the feasibility of this method, we diagnosed a very near field plume of a 15 kW AF-MPDT and obtained the distribution of parameters with a spatial resolution of 0.23 mm. The results show that the maximum and are located on the centerline near the exit plane of the thruster and decrease with both radial and axial distances. The second peak appears in the radial direction of both the and distributions, while their zones do not overlap. A comparative analysis of measurement results obtained by this method and the Langmuir probe method is performed to prove its accuracy. Moreover, this measurement method has the advantages of high spatial resolution and the ability to diagnose high-density and magnetized plasma. This method can be also applied to other optically thin axisymmetric low-temperature argon plasmas.

Flight Tests of Passive Flow Control for Suppression of Cavity Aeroacoustics

This research focused on furthering the current understanding of the flowfield inside a cavity under typical flight conditions. The flight envelope included altitudes of up to 40,000 ft and maximum Mach number of 1.28 and included several test points between Mach 0.9 and 1.1, which are challenging to isolate in wind tunnel testing due to tunnel wall interference. The pod was instrumented to collect aeroacoustic pressures and qualitative flowfield imaging for a cavity with a length-to-depth ratio of four. Test missions were flown with several different passive flow control configurations, and their respective effect on acoustic properties and flow patterns were collected. Flow control comprising a rod-in-a-crossflow or a traditional sawtooth spoiler was generally found to reduce sound pressure level, though the extent of effectiveness varied with flight condition. The flow control devices installed at or below the cavity exit plane were found to have little influence. Tuft-based flow visualization complemented the effort, for one example, by demonstrating the outward flow near the cavity lip for the sawtooth spoiler.

Flowfield of a helicopter submerged inlet with power output shaft

The submerged inlet is an attractive configuration for advanced helicopters due to its high stealth performance and low external drag. In this paper, a submerged inlet, integrated with a ROBIN helicopter fuselage and a simplified power output shaft, is experimentally and numerically investigated to obtain the basic flow characteristics under a freestream velocity of 23.6 m/s. The results indicate that the pylon ahead of the inlet induces a horseshoe vortex. Though the vortex is ingested into the inlet, it has little effect on the internal flows and can be neglected. When the airflow enters into the inlet, it interacts with the shaft with a large incidence angle, yielding a vortex pair. At the leeside of the shaft, the two side flows of the shaft impinge at the center plane, generating a local high-pressure region at the azimuthal angle of 180°, which forces the boundary layer to roll up a counter-rotating vortex pair. In addition, the airflow adjacent to the cowl lip accelerates rapidly, resulting in a local low-pressure region at the azimuthal angle of 0°. Therefore, the inlet duct has a strong circumferential pressure gradient, which originates from an azimuthal angle of 180° to 0° and induces a vortex pair at the azimuthal angle of 0°. The three vortex pairs are the main origins of the distortion at the duct exit plane, among which the one near the cowl lip with the azimuthal angle of 0° plays the dominant role. Additionally, as the velocity ratio increases from 3.9 to 5.5, the circumferential pressure gradient and the cowl lip vortex get intensified, which causes that the total-pressure recovery coefficient drops by 0.5% and the distortion index increases by 28%. A submerged inlet, integrated with a ROBIN helicopter fuselage and a simplified power output shaft, is experimentally and numerically investigated. Three vortex pairs, which locate at the azimuthal angle of 0°, the leeside of the shaft, and 180° of the inlet surface, are the main origins of the distortion of the inlet, among which the one near the cowl lip with the azimuthal angle of 0° plays the dominant role. As the velocity ratio increases, the circumferential pressure gradient gets intensified, leading to stronger vortex pairs.

Computational estimation of point-spread-function field in axially symmetric optical systems

The Point-Spread-Function (PSF) field characterizes the geometric and diffractive properties of most optical system designs for imaging applications. The conventional method of computing the PSF field is carried out in two steps: first, geometrical computation of the wave aberration function, and then, diffraction evaluation. That method usually applies two approximations: (1) the exit and entrance pupil ray intersection coordinates are equal; and (2) the optical field amplitude distribution does not change between exit and entrance pupil planes. We propose here a variation of the conventional method that allows overcoming these two approximations in axially symmetric optical systems. To show the potentials of the method we applied it to a Cooke triplet. We analyse the influence of the number of traced rays on the measured accuracy of the method. We also evaluate the differences in the PSF estimation accuracy with respect to a method where the two aforementioned approximations were applied.

Nozzle design influence on the steam-driven ejector

In the present paper, a series of numerical simulations of wet steam flows within ejectors distinguished by four primary nozzle contours have been carried out with the objective to evaluate the overall ejector performance. The studied primary nozzle contours are the following: a standard Laval nozzle (LAVAL); a nozzle that is designed in order to provide Constant Expansion Rate (CER); and two nozzles (MOC_SHORT, MOC_LONG) that are designed by employing axisymmetric Method of Characteristics. At first, the wet steam flow simulation throughout solely nozzles are carried out. Within the CFD simulations routines most valuable flow parameters are compared: expansion rates, boundary layer displacement and momentum thicknesses; entropy generation rate from various thermodynamic forces; thrust and liquid mass fraction across the nozzle exit plane. A comprehensive analysis of solely nozzles revealed that the CER nozzle contour is the most aerodynamically efficient design, which provides the minimum liquid mass fraction by the nozzle exit and possess the minimum entropy generation rate. The CFD results analysis of a full ejector domain revealed that an ejector with a MOC_SHORT primary nozzle contour provides the maximum performance in terms of secondary mass flow rate. At that, the secondary mass flow rate in MOC_SHORT nozzle contour case is almost 4% greater than for the CER nozzle design case.

Impact of fuel hydrogen content on non-volatile particulate matter emitted from an aircraft auxiliary power unit measured with standardised reference systems

Replacement of conventional petroleum jet fuel with sustainable aviation fuels (SAFs) can significantly reduce non-volatile Particulate Matter (nvPM) emissions from aircraft main engines and auxiliary power units (APUs). As part of the Initiative Towards sustAinable Kerosene for Aviation (ITAKA) project, the impact of fuel hydrogen content on nvPM number and mass emissions and particle size distributions were investigated using a GTCP85 APU burning blends of conventional (Jet A-1) and Hydrotreated Esters and Fatty Acids (HEFA)-derived (Used Cooking Oil and Camelina) aviation fuels. The measurements were conducted during two separate test campaigns performed three years apart, each employing a different regulatory compliant sampling and measurement reference system for aircraft engine nvPM emissions. The objective was to investigate the correlation of fuel hydrogen content with nvPM number and mass emissions at the engine exit plane (EEP) independent of fuel composition, measurement system, and ambient conditions. The nvPM number and mass emissions and size distributions systematically decreased with increasing fuel hydrogen content regardless of the fuel composition or APU operating condition. The measured nvPM emissions were particle loss-corrected to the EEP and normalised to a common fuel hydrogen content. Similar rates of nvPM reductions were observed for both test campaigns at all investigated APU operating conditions, confirming that engine exit nvPM reductions correlate with fuel hydrogen content for fuels of relatively similar compositions. This analysis method can be applied to emissions data from other engine types to compare the reduction in nvPM emissions for sustainable aviation fuels and blends.

Phase transformation of a vortex beam in a liquid-based inferior mirage

We study how a liquid-based inferior mirage, obtained by layering distilled water and ethanol, transforms the phase structure of a light beam possessing a helical wave front. An inferior mirage amounts for one total internal reflection, which effectively reverses the handedness of the wave front. We show that this transformation is accompanied by smooth unidirectional astigmatic changes and variations of the non-canonical strength of the phase singularity nested in the beam. A skew in the beam intensity distribution is observed where the phase singularity is inverted and allows the direct measurement of the topological charge of the beam. Freely propagating, partially inverted beams possessing spatially varying orbital fluxes can be obtained at the exit plane of the solution by adjusting the incidence conditions of the beam. This work lays the foundations for phase engineering of light beams in liquid-based optical mirages.

Renormalisation of particle distributions in an initially-biased turbulent jet by swirl and radial injection

The distribution of particles in a jet with an initial bias is assessed using planar nephelometry, which is a planar imaging technique using Mie-scattering. Initial biases in particle distributions can have large and significant effects, particularly in solid fuel combustion systems, where inhomogeneities will affect the local and global performance of a flame. In this study, a two-phase jet flow with a controlled initial bias is generated at a Stokes number, based on the large-eddy time-scale, of SkD=0.39, and various strategies are explored to correct this bias. The investigation examines the usefulness of injecting air at various momentum ratios and through various configurations upstream from the nozzle exit plane. The study examines the differential roles of swirl (tangential) and radial injection of air at various locations from the flow exit. Radial injection of momentum is found to be superior to injection of swirl, consistent with swirl being a magnifier of eccentricities of a flow. Radial injection upstream from the nozzle can achieve a re-homogenised flow to develop prior to the exit, at least for the present Stokes numbers, but can also lead to downstream biases in particle distribution for some conditions. Injecting momentum directly at the nozzle exit does not allow a symmetrical distribution to be achieved, but does have an observed influence on the emerging flow-field.

Influence of Geometric Parameters on Overall Performance of High Bypass Ratio Turbofan Nacelle and Exhaust System

In order to understand the coupled effect of the nacelle and exhaust system and to improve their overall performance, we studied the aerodynamic performance and the flow characteristics of the high bypass ratio turbofan nacelle and exhaust system by numerical simulation. The geometric parameters of a nacelle and exhaust system (e.g., the contraction ratio of the cowl afterbody and the fan nozzle exit angle) were investigated to evaluate their influence on the overall performance of the nacelle and exhaust system. The related flow mechanism was explored as well. The results show that the flow field of the nacelle and exhaust system under the mid-cruise condition exhibits characteristics of transonic flow. A stagnation zone exits at the nacelle lip and there is a velocity peak at the nacelle forebody. There exist a number of complex flow phenomena (such as shockwave, expansion wave, shear flow and shock wave-boundary layer interaction) in the downstream of the fan nozzle exit plane. The magnitude of the fan nozzle thrust or the intake ram drag is much higher than that of the additional drag, the nacelle drag or the core nozzle thrust. And for the nacelle drag, the friction drag of the cowl is in the same order of magnitude as the pressure drag of the cowl, the core cowl and the plug. But it is much larger than the friction drag of the core cowl and the plug. The effective thrust increases by 4.7% as the contraction ratio of the cowl afterbody increases; and it increases by 2.4% as the fan nozzle exit angle increases. The expansion degree of the fanjet flow, the shock wave strength and location, and the existence of the flow separation or second shock wave are influenced by the contraction ratio of the cowl afterbody and the fan nozzle exit angle. These phenomena have effects on the pressure distribution of the core cowl and the surrounding fanjet flow velocity, and hence they further affect the nacelle drag. The increase in the fan nozzle exit angle can noticeably reduce the thrust of the fan nozzle.

Numerical simulation of a methane-oxygen rotating detonation rocket engine

The rotating detonation engine (RDE) is an important realization of pressure gain combustion for rocket applications. The RDE system is characterized by a highly unsteady flow field, with multiple reflected pressure waves following detonation and an entrainment of partially-burnt gases in the post-detonation region. While experimental efforts have provided macroscopic properties of RDE operation, limited accessibility for optical and flow-field diagnostic equipment constrain the understanding of mechanisms that lend to wave stability, controllability, and sustainability. To this end, high-fidelity numerical simulations of a methane-oxygen rotating detonation rocket engine (RDRE) with an impinging discrete injection scheme are performed to provide detailed insight into the detonation and mixing physics and anomalous behavior within the system. Two primary detonation waves reside at a standoff distance from the base of the channel, with peak detonation heat release at approximately 10 mm from the injection plane. The high plenum pressures and micro-nozzle injector geometry contribute to fairly stiff injectors that are minimally affected by the passing detonation wave. There is no large scale circulation observed in the reactant mixing region, and the fuel distribution is asymmetric with a rich mixture attached to the inner wall of the annulus. The detonation waves’ strengths spatially fluctuate, with large variations in local wave speed and flow compression. The flow field is characterized by parasitic combustion of the fresh reactant mixture as well as post-detonation deflagration of residual gases. By the exit plane of the RDRE, approximately 95.7% of the fuel has been consumed. In this work, a detailed statistical analysis of the interaction between mixing and detonation is presented. The results highlight the merit of high-fidelity numerical studies in investigating an RDRE system and the outcomes may be used to improve its performance.

Experimental and numerical investigation of expansion corner effects on isolator performance

Abstract In the present study, the effect of expansion corner on suddenly expanded flow process has been studied. Experimental investigations have been carried out on a convergent-divergent (C-D) nozzle and isolator duct, where the expansion of the channel is formed through the presence of a 1, 2 and 3 expansion corners (EC) respectively. Flow from nozzle exit of the nozzle of Mach, M = 2.0 was suddenly expanded into the axi-symmetric duct having a cross sectional area of 4.84 times the nozzle exit area. The wall static pressure along the length of the duct and the Pitot pressure at the exit plane of the duct were measured for all the configurations. Computational fluid dynamics (CFD) technique was employed for visualizing the shock-train in the expanded duct. The isolator with one expansion corner was found to be more efficient in achieving a high static pressure rise. The experimental and numerical wall static pressure distribution values were compared for isolators with EC = 2 and found to be in good agreement with each other with a maximum absolute percentage deviation of 11%.

Experimental and numerical investigation of expansion corner effects on isolator performance

Abstract In the present study, the effect of expansion corner on suddenly expanded flow process has been studied. Experimental investigations have been carried out on a convergent-divergent (C-D) nozzle and isolator duct, where the expansion of the channel is formed through the presence of a 1, 2 and 3 expansion corners (EC) respectively. Flow from nozzle exit of the nozzle of Mach, M = 2.0 was suddenly expanded into the axi-symmetric duct having a cross sectional area of 4.84 times the nozzle exit area. The wall static pressure along the length of the duct and the Pitot pressure at the exit plane of the duct were measured for all the configurations. Computational fluid dynamics (CFD) technique was employed for visualizing the shock-train in the expanded duct. The isolator with one expansion corner was found to be more efficient in achieving a high static pressure rise. The experimental and numerical wall static pressure distribution values were compared for isolators with EC = 2 and found to be in good agreement with each other with a maximum absolute percentage deviation of 11%.

Release characteristics of overpressurised gas from complex vents: implications for volcanic hazards

Many explosive volcanic eruptions produce underexpanded starting gas-particle jets. The dynamics of the accompanying pyroclast ejection can be affected by several parameters, including magma texture, gas overpressure, erupted volume and geometry. With respect to the latter, volcanic craters and vents are often highly asymmetrical. Here, we experimentally evaluate the effect of vent asymmetry on gas expansion behaviour and gas jet dynamics directly above the vent. The vent geometries chosen for this study are based on field observations. The novel element of the vent geometry investigated herein is an inclined exit plane (5, 15, 30° slant angle) in combination with cylindrical and diverging inner geometries. In a vertical setup, these modifications yield both laterally variable spreading angles as well as a diversion of the jets, where inner geometry (cylindrical/diverging) controls the direction of the inclination. Both the spreading angle and the inclination of the jet are highly sensitive to reservoir (conduit) pressure and slant angle. Increasing starting reservoir pressure and slant angle yield (1) a maximum spreading angle (up to 62°) and (2) a maximum jet inclination for cylindrical vents (up to 13°). Our experiments thus constrain geometric contributions to the mechanisms controlling eruption jet dynamics with implications for the generation of asymmetrical distributions of proximal hazards around volcanic vents.

Plasmas generated by electron beam passing through diamond window

In this paper, we report the generation of Ar plasmas by using an electron beam passing through a diamond window. It is observed that plasma brightness varies with the electron beam energy and the duty cycle. The transmission properties of the electrons in a 5-μm-thick diamond film are investigated by both experiments and Monte Carlo simulation. The measured transmittance of the electron beam increases from 20% to 80% by increasing the incident energy of the electron beam on the diamond window from 32.5 keV to 50 keV. The diamond window was checked using a CCD camera and SEM after plasma generation experiments. It is found that the focused electron beam passes through the central region of the diamond window, from the incident plane to the exit plane, and the cross section of the diamond window gradually changes from textured morphology to a featureless structure, indicating that the main energy loss of electrons in the diamond window occurs in the later stage of the transmitting journey. The interaction between the electron and the diamond causes the structure change of the diamond. The simulated transmittances of the electron beam with respect to its incident energy are in agreement with the experimental results.

Experimental Assessment of Corrugated Rectangular Actuators on Supersonic Jet Mixing

To improve the stealth capability of a military aircraft, the reduction in core length is essential to reduce the heat signature and the noise characteristics of the engine exhaust. The efficacy of rectangular vortex generators in achieving these objectives has been demonstrated by several researchers, owing to their simplicity. One way of producing the mixed-size vortices is by providing corrugations on the edge of the tab (actuator). Therefore, in the current study, two tabs of aspect ratio 1.5, mounted diametrically opposite to each other at the outlet of a Mach 1.73 circular nozzle, are examined at varying levels of expansions, ranging from overexpanded to underexpanded jet states. In addition, to generate the mixed-size vortices, three corrugation geometries, i.e., rectangular, triangular, and semicircular, are configured along the tab edges. Both quantitative and qualitative investigations are carried out by using the pitot probe to measure the stagnation pressures and by utilizing a shadowgraph technique to visualize the flow field. The corrugated tabs generated a significant mixing, and among them, the tabs with triangular corrugations are found to be most effective. A maximum reduction of about 99.7% in the supersonic core is obtained with triangular corrugated tabs at near-correct-expansion, corresponding to nozzle pressure ratio (NPR) 5. Interestingly, the semicircular corrugated tab significantly reduces the asymmetry near the nozzle exit plane. The shadowgraph images confirm the efficacy of different corrugated tabs in reducing the strength of the waves, prevalent in the supersonic core.

Development of the impulse and thrust for laminar starting jets with finite discharged volume

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Acoustic feedback loops for screech tones of underexpanded free round jets at different modes

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Film cooling performance and flow field of compounded double jet holes with trench

Film cooling performance of a novel double jet configuration is numerically analysed. The proposed scheme has a primary compound angle hole along with a secondary hole compounded in the laterally opposite direction. The primary hole is provided with a trench at the coolant hole exit. Numerical analysis is performed using RANS equations along with standard k-ω turbulence model. The numerical scheme is validated through the excellent agreement between the available experimental data for tangential, compound, and crater designs having similar primary hole geometric parameters. It is seen that the proposed compounded double jet trench design (DJT) combines the advantages of compound angle, double jet and trenched configurations and provides better lateral and uniform coolant distribution, reduced jet lift off, and higher coolant attachment in the stream wise direction. The vortex mechanisms responsible for the coolant spread in near-hole and down stream regions are identified. The various flow structures that are responsible for higher effectiveness are compared with other configurations. The effect of free stream turbulence at various blowing ratios is studied and found that the increase of turbulence produced significant increase in lateral effectiveness at higher blowing ratios. This effect is predominant at far downstream locations of the coolant exit plane.