Mixing Duct Research Articles

Linear analysis of a swirling jet with a realistic swirler model

The dynamics of an axisymmetrical swirling jet is studied via global linear stability and resolvent analyses. The modeled flow represents a combustor-like swirling jet, that is turbulent, compressible, non-parallel, and enclosed. In particular, the computational domain embeds a realistic axisymmetrical swirler model to resolve the mode conversion process. Swirl fluctuations are non-negligible on this configuration representative of a swirl burner, and match the analytical mode shapes of inertial waves of an inviscid uniform flow as obtained from global stability analysis. The stability map presents two eigenvalues driving a modal amplification. These eigenmodes couple a standing acoustic wave sustained in the mixing duct and the combustion chamber with the Kelvin-Helmholtz mechanism at the mixing duct exit and the acoustic-vorticity mode conversion process at the swirler, and act as a frequency selection criterion. Finally, the most amplified forcing from the resolvent analysis is similar to an unsteady heat source in the combustion chamber, and the identified optimal amplification mechanism is likely to be triggered in reacting flow with unsteady heat release rate.

Mechanisms of lobed jet mixing: Comparison of bilinear, rectangular, and circular lobed mixers

In this study, bilinear lobed mixers (BLMs), rectangular lobed mixers (RLMs), and circular lobed mixers (CLMs), with two sectional area ratios of the bypass channel to the core channel, were configured while maintaining the same lobe angle, lobe height, lobed nozzle length, mixing duct length, and core sectional area. Using three-dimensional steady Reynolds-averaged Navier–Stokes equations and the SST k–ω turbulence model, numerical simulations of jet mixing in these lobed mixers at three velocity ratios of the bypass stream to the core stream were performed. The jet mixing characteristics and performances of the three types of lobed mixers were analyzed. Through a comprehensive and in-depth analysis of the jet-mixing flow fields of the three types of lobed mixers, more universal mechanisms of lobed jet mixing were revealed. Systematic and meaningful conclusions were drawn from four aspects: the formation and evolution mechanisms of the cross flow, longitudinal vortices, heat and mass convection, and orthogonal vortex strip. The findings from this study hold significant implications for the development of lobed mixers and exploration of lobed jet-mixing mechanisms.

Research on the Infrared Radiation Suppression of the High-Temperature Components of the Helicopter with an Integrated Infrared Suppressor

The integrated infrared suppressor can reduce the infrared radiation signal of a helicopter and is compatible with radar-acoustic stealth. However, the issues that are caused by the integrated infrared suppressor, such as temperature increases on the rear fuselage surface and a lack of shielding at the exhaust port, need to be addressed, in order to further improve the infrared stealth capability of the helicopter. Aiming at this, the effects of the ambient temperature, fuselage surface emissivity, mixing duct shielding, and exhaust port shielding on the infrared radiation characteristics of the helicopter are studied with numerical simulation. The results show that the infrared radiation intensity of the helicopter, in 3–5 μm band and 8–14 μm band, decreases by about 20% and 10%, respectively, for every 6 K reduction in the ambient temperature. As the emissivity of the rear fuselage surface reduces from 0.8 to 0.5, the helicopter’s infrared radiation intensity, in a 3–5 μm band and a 8–14 μm band, decreases by about 6% and by about 4% and 1.3%, respectively, after the mixing duct is equipped with a shielding sheath. Installing deflectors at the exhaust port of the fuselage can prevent the detection rays from detecting the high-temperature components inside the fuselage, and when the emissivity of the deflectors is reduced from 0.8 to less than 0.5, or the deflectors are cooled by more than 80 K, they begin to play a role in suppressing the infrared radiation at the bottom of the helicopter.

Low-frequency unsteadiness of recompression shock structures in the diffuser of supersonic ejectors

Supersonic ejectors are passive gasdynamic devices that compress a low-pressure fluid by utilizing the kinetic energy of a high-pressure fluid in a variable area duct. The ejector consists of a primary supersonic nozzle in a mixing duct where the secondary flow is entrained and mixed. The mixed flow can undergo a series of recompression shocks resulting in a subsonic flow in the diverging portion to aid pressure recovery. Recompression shocks usually lead to unsteady shock boundary layer interactions. The performance of the ejector is influenced by shear layers, shock and expansion waves, and their mutual interactions. While existing literature has extensively dealt with mixing of the primary and secondary flows, the unsteadiness in flow resulting from recompression shocks has been seldom investigated. Fluctuations in pressure due to the unsteadiness of the shock often lead structural fatigue issues. This paper reports a detailed investigation on low-frequency unsteadiness of recompression shock using high-speed schlieren images and dynamic pressure measurements. Modal analyses using proper orthogonal decomposition and dynamic mode decomposition techniques are used to determine the dominant spatial modes and associated frequencies. Multimodal frequencies ranging between 80 and 300 Hz are observed. These findings are further corroborated by Fourier and wavelet transformations of the experimentally measured wall static pressure signals. Subsequently, scaling parameter is established for the dominant frequencies based on flow velocities upstream of the shock and the distance between two consecutive shocks. This results in a unique scaling frequency of 4.58% ± 18%, for the recompression shock independent of operating conditions.

Artificial neural network based shape optimization of supersonic ejectors in the critical flow regime

Supersonic ejectors are widely used in energy, refrigeration, and aerospace propulsion applications. A supersonic ejector is a passive gasdynamic device consisting of a converging–diverging nozzle located in a varying area mixing duct where the secondary fluid is compressed by mixing with high momentum primary fluid. The entrainment ratio (ER, ratio of mass flow rates of the secondary to primary flow) is dependent on the mixing duct geometry, necessitating shape optimization. Complex gasdynamic interactions in the ejector are not fully understood, thus making physics-based or CFD modeling, a challenging exercise. In this paper, experimental data for ejector performance with varying geometries and operating parameters, viz. area ratio, primary flow Mach number, mixing duct length to height ratio (L/D), and stagnation pressure ratio is presented. The mixing length parameter is quantified using data derived from high-speed schlieren images. Improvement in ER to the extent of 20% is achieved by changing L/D to 15 for the AR of 3.33, with further increase in L/D resulting in a reduction in ER. Subsequently, computationally inexpensive model using ANN is developed to predict the ejector performance using the experimental database. The model predicts ER with an accuracy of 7%. Finally, the ANN model is used to optimize the ejector geometry for critical mode operation, which provides a 30% enhancement in entrainment ratio compared to the baseline geometry.

An Experimental Study Of Natural Gas Jet-Flames In Crossflow With High Hydrogen Admixture At Elevated Pressure Conditions

A series of high hydrogen-content jet flames in crossflow are investigated experimentally using high-speed (10 kHz) stereoscopic particle image velocimetry, dual-plane laser-induced fluorescence of the hydroxyl radical, and OH* chemiluminescence imaging. The flames were fueled with natural gas enriched to 70-, 80-, 90- and 100% H2, by volume at conditions approaching those found in the mixing duct of a modern gas turbine engine, specifically in close confinement, at 7 bars pressure and with a crossflow preheat of 511 K. The measurements indicate the recirculation zone on the leeward side of the jet plays a critical role in stabilization of the lifted jet flame in crossflow. The measurements show that local heat-release from a flame in this region acts as a self-reinforcing feedback mechanism, strengthening the recirculation zone and increasing local residence time there. This, in turn, creates conditions more favorable to flame stabilization. The measurements also showed how close confinement of the JICF by the flow channel walls affects jet-penetration and growth of the counter-rotating vortex pair.

Modeling the Transport of Fuel Mixture Perturbations and Entropy Waves in the Linearized Framework

Abstract In this paper, a new method is introduced to model the transport of entropy waves and equivalence ratio fluctuations in turbulent flows. The model is based on the Navier–Stokes equations and includes a transport equation for a passive scalar, which may stand for entropy or equivalence ratio fluctuations. The equations are linearized around the mean turbulent fields. These serve as the input to the model in addition to a turbulent eddy viscosity, which accounts for turbulent diffusion of the perturbations. Based on these inputs, the framework is able to predict the linear response of the flow velocity and passive scalar to harmonic perturbations that are imposed at the boundaries of the computational domain. These, in this study, are fluctuations in the passive scalar and/or velocities at the inlet of a channel flow. The code is first validated against analytic results, showing very good agreement. Then, the method is applied to predict the convection, mean flow dispersion, and turbulent mixing of passive scalar fluctuations in a turbulent channel flow, which has been studied in previous work with direct numerical simulations (DNS). Results show that our code reproduces the dynamics of coherent passive scalar transport in the DNS with very high accuracy and low numerical costs. Furthermore, we demonstrate that turbulent mixing has a significant effect on the transport of the passive scalar fluctuations. Finally, we apply the method to explain experimental observations of transport of equivalence ratio fluctuations in the mixing duct of a model burner.

Research on the influence of integrated infrared suppressor exhaust angle on exhaust plume and helicopter infrared radiation

A numerical investigation is performed in the current study to illustrate the influences of exhaust angle of the integrated infrared suppressor on plume flow and helicopter infrared radiation, under the hover and cruise statuses. Three exhaust models are concerned, including the exhaust direction at 60°, 45°, and 18° angles to the horizontal plane. The results show that the heating effect of the exhaust plume on the fuselage is modified with the decrease of the exhaust angle. However, the decrease in exhaust angle will reduce the pumping coefficient of the mixing duct to some extent, resulting in an increase of exhaust temperature. For the model with an exhaust angle of 18°, three sets of comparative tests were designed. The results show that when there is no airflow outlet at the bottom of the fuselage, the surface temperature of the fuselage will increase significantly, and the position of the airflow outlet at the bottom of the fuselage also has a certain influence on the temperature field of the fuselage. On the horizontal detection plane, the infrared radiation intensity of the model with an exhaust angle of 60° and 45° is significantly lower than that with an exhaust angle of 18°. In the vertical detection plane, especially in the direction below the fuselage, the infrared radiation intensity of the model with an exhaust angle of 18° is the smallest. When compared to the hover status, the cruise status reduces the infrared radiation intensity somewhat.

Interaction of equivalence ratio fluctuations and flow fluctuations in acoustically forced swirl flames

The generation and turbulent transport of temporal equivalence ratio fluctuations in a swirl combustor are experimentally investigated and compared to a one-dimensional transport model. These fluctuations are generated by acoustic perturbations at the fuel injector and play a crucial role in the feedback loop leading to thermoacoustic instabilities. The focus of this investigation lies on the interplay between fuel fluctuations and coherent vortical structures that are both affected by the acoustic forcing. To this end, optical diagnostics are applied inside the mixing duct and in the combustion chamber, housing a turbulent swirl flame. The flame was acoustically perturbed to obtain phase-averaged spatially resolved flow and equivalence ratio fluctuations, which allow the determination of flux-based local and global mixing transfer functions. Measurements show that the mode-conversion model that predicts the generation of equivalence ratio fluctuations at the injector holds for linear acoustic forcing amplitudes, but it fails for non-linear amplitudes. The global (radially integrated) transport of fuel fluctuations from the injector to the flame is reasonably well approximated by a one-dimensional transport model with an effective diffusivity that accounts for turbulent diffusion and dispersion. This approach however, fails to recover critical details of the mixing transfer function, which is caused by non-local interaction of flow and fuel fluctuations. This effect becomes even more pronounced for non-linear forcing amplitudes where strong coherent fluctuations induce a non-trivial frequency dependence of the mixing process. The mechanisms resolved in this study suggest that non-local interference of fuel fluctuations and coherent flow fluctuations is significant for the transport of global equivalence ratio fluctuations at linear acoustic amplitudes and crucial for non-linear amplitudes. To improve future predictions and facilitate a satisfactory modelling, a non-local, two-dimensional approach is necessary.

Sea-Level Static Tests of Rocket–Ramjet Combined Cycle Engine Model

A rocket–ramjet combined cycle engine model, embedding twin rocket chambers driven with gaseous hydrogen and oxygen on the top wall side of a scramjet flowpath, was tested in its ejector-mode operation under sea-level static conditions. Gaseous hydrogen was also injected through secondary injector orifices to pressurize a ramjet combustor located downstream of the rocket chambers. Mixing between the hot rocket plume and cold airflow as well as combustion of residual fuel within the plume with the airflow caused entropy and static pressure increases in the constant-area mixing duct in the original flowpath design, resulting in a limited airflow rate. The mixing duct was redesigned to incorporate divergence from its onset to counter the pressure rise. With the diverging flowpath configuration, the airflow rate was increased by 40 percent. However, this modified flowpath geometry resulted in the formation of a low-speed area, through which the pressure rise in the ramjet combustor penetrated the mixing duct and reduced the incoming airflow rate. Without mechanical contraction near the engine exit, increasing pressure in the diverging portion of the flowpath was difficult, which in turn, resulted in poor mixing between the airflow and the rocket plume. Balancing these factors and additional mixing enhancement are required. The engine performance is then summarized.

Effects of Hydrogen-Enrichment on Flame-Holding of Natural Gas Jet Flames in Crossflow at Elevated Temperature and Pressure

The effect of hydrogen (mathrm {H}_{mathrm {2}}) enrichment on the flame-holding characteristics of two natural gas jet flames in crossflow is investigated here, experimentally. The flame and flowfield measurements are analyzed using simultaneously acquired high-speed (10 kHz) stereoscopic particle image velocimetry, planar laser-induced fluorescence of the hydroxyl radical, and OH* chemiluminescence. The flames, enriched with 20% and 40% mathrm {H}_{mathrm {2}}, by volume, are studied at conditions typical of the mixing duct of a modern gas turbine engine; specifically in confinement, at 10 bars, and with a crossflow preheat of 530 K. Consistent with previous findings, the 40% mathrm {H}_{mathrm {2}} flame was found to be stabilized on the windward and leeward side of the jet, while the 20% mathrm {H}_{mathrm {2}} flame was stabilized only on the leeward side. Analysis of mean and instantaneous velocity fields showed no major differences in the trajectories and principal compressive strain fields of the two flames. The presence of the windward stabilized flame in the 40% mathrm {H}_{mathrm {2}} case was, however, found to decrease the centerline velocity decay and greatly reduce or eliminate large scale vortices along the windward shear layer. The difference in the flame-holding here was attributed to the difference in the extinction strain rate from the addition of hydrogen, which would impact the local and global extinction of the flame along the high shear windward region of the flame.

Numerical investigation on flow and mixing characteristics inside a converging-diverging mixing duct of rocket-based combined-cycle engine in ejector mode

The supersonic mixing layers with complex background waves are formed by the rocket jet and the entrained air in the mixing section of rocket-based combined-cycle (RBCC) engine during ejector mode. To achieve mixing enhancement of primary and secondary flows in confined space, a converging-diverging mixing duct was designed in this study. For a typical flight regime of RBCC ejector mode, the effects of the mixer geometry including the contraction ratio, the throat position of mixer and the converging angle on the flow structures and mixing characteristics in the converging-diverging mixing duct under no backpressure condition were numerically investigated. A detailed discussion on the mechanism of mixing enhancement in this mixer configuration was also presented. The results indicate that appropriately increasing the contraction ratio can significantly strengthen the interactions of background shock waves with supersonic mixing layer in the converging-diverging mixing duct, which effectively promotes the rapid and sufficient mixing between primary flow and secondary flow. Besides, shortening the length from the mixer inlet to the throat position can enhance the growth rate of mixing layer and the mixing efficiency of two flows in the upstream region of converging-diverging mixing duct to some extent. However, in the downstream far-field region, the variations of throat position have little impact on the scalar mixing process in supersonic mixing layer. Moreover, it is found that increasing the converging angle has no obvious effect on intensifying the mixing performance of the converging-diverging mixing duct. On the contrary, it may lead to a dramatic decrease in total pressure recovery. Further analysis reveals that the oblique shocks generated at the compression corner and the separation shocks formed in the divergent section interact with the supersonic mixing layer, which contributes to mixing enhancement in the converging-diverging mixing duct.

Experimental investigation of shock train unsteady movement in a mixing duct

The structure and dynamic of the shock train in a mixing duct are investigated experimentally in a supersonic air breathing wind tunnel. High-speed schlieren technique and pressure measurements are utilized for the data acquisition. Nozzle clapboards are applied to separate the incoming flow into upper secondary flow, primary flow, and lower secondary flow. The Mach number of all incoming flows is 2.05, whereas the static pressures between primary and secondary flow stay unmatched. Experimental results indicate that nozzle clapboards lead to complex background waves and mixing layers in the mixing duct. As the throttling valve rotates at a constant speed, two motion patterns alternatively govern the shock train movement, namely jump movement and slow movement. Judging from the path of shock train leading edges, adverse pressure gradient at the wall of background flow induces jump movement, while the slow movement only appears at the zone of favorable pressure gradient. Besides, leading shock structures continuously change in the forward propagation process. Four types of leading shock structures are discovered namely regular reflection, convex-stem Mach reflection, straight-stem Mach reflection, and concave-stem Mach reflection. Based on the numerical results of the background flow field, these special structures are explained theoretically by analyzing the variation of the wavefront Ma contour lines.

Mechanisms of lobed jet mixing: About circularly alternating-lobe mixers

Two configurations of circularly arranged alternating-lobe nozzles were adopted to form lobed mixers with/without a mixing duct. The jet mixing of each mixer was numerically simulated with the unchanged initial conditions of the primary and secondary streams, except the altered initial velocity of the secondary stream. The jet-mixing mechanisms of the circularly alternating-lobe mixers were synthetically analysed by combining the evolution of the flow field structures and the process of heat and mass transfer in the mixing field. It is found that the transverse flow is usually caused by the lobed geometry, and the entrainment of the primary stream also plays a role in certain circumstances. There are two mechanisms for the deflection of the transverse flow at the lobe peaks and troughs. One of these mechanisms is to be suppressed to deflect the flow while the other is deflected by the reaction force and induction effect. The primary and secondary streams deflect to bring the transverse interval between them, and subsequently, the streamwise vortex core appears at the transverse interval. The deflected flow consistently “digging” in the radial and circumferential radiation increases the dimension of the streamwise vortices. The transverse flow velocity decreases and the direction becomes unstable leading to the breakdown of the streamwise vortices. The transverse flow brings the heat and mass transfer. Under the two mechanisms, the frontiers of the primary and secondary streams deflect. Initially, the primary and secondary streams flow around the streamwise vortex core. Subsequently, the mixed stream flows around the vortex core, and the mixing stream area gradually expands outward. The heat and mass transfer decrease in scale when the streamwise vortices break down. Because of the velocity gradient at the interface, shear instability occurs to generate the normal vortex ring. The heat and mass transfer pushes the interface, which leads to the stretch of the normal vortex ring. The mixing speed varies due to the heat and mass transfer. The velocity gradient decreases fast in the rapid mixing segment, where the normal vortex ring breaks first.

Effects of component geometries and inflow conditions on ejector operational mode

The ejector system is a useful device for creating high altitude conditions for ground tests of supersonic engines. The ejector performance can be immediately changed by varying the ejector operational mode. The ejector nozzle pressure ratio is well known to have a significant effect of the operational mode of an ejector. However, the effects of the mixing duct length and other geometric design parameters on the ejector mode change have not been clearly determined. In this study, the effects of ejector component geometries and inflow conditions on ejector operational mode are investigated by numerical analysis. By changing the inflow conditions and geometric parameters, twelve test cases are studied. Using the numerical test results, the flow pattern and suction pressure performance of the ejector with a fixed secondary mass flow rate are compared. In the numerical test results, a high primary nozzle stagnation pressure induces a highly underexpanded flow, resulting in the critical operational mode. For the critical operational mode, the mixing duct must be sufficiently long to accommodate the shock train and the choking zone. Primary nozzles with wide angles also induce widely expanded nozzle flows and result in the critical operational mode.

Suggestions on investigations of lobed jet mixing

Jet mixings of different constructional lobed mixers comprised of lobed nozzles with and without a mixing duct were investigated numerically. Based on the heat and mass transfer investigation in the streamwise vortices and the “stretch” and “cut” of the normal vortex ring, suggestions on investigations of lobed jet mixing were provided. A more reasonable mechanism for lobed jet mixing was suggested to be revealed when combined fluid dynamics and heat and mass transfer were being investigated. Differences in fluid dynamics and heat and mass transfer with different constructional lobed mixers meant different mechanisms.

Numerical investigation of shock train unsteady movement in a mixing duct

Supersonic mixing layers and complex background waves exist in the mixing duct of the RBCC (Rocket-Based Combined Cycle) engine and supersonic-supersonic ejector. Different from the isolator with uniform incoming flow, the shock train in a mixing duct has special structures and motion characteristics. To study the characteristics of shock train forward movement in a mixing duct, a 2d unsteady numerical simulation was conducted. The results indicated that many unsteadiness factors existed in the background flow field such as the development of mixing layers, and multiple complex interactions. With backpressure linearly increasing over flow time, the shock train moved upstream in the process of which two forward movement patterns were found, namely slow pattern and jump pattern. Two patterns alternatively governed the whole forward movement, resulting in periodical changes of leading shock structures. The leading shocks were ‘stretched’ and ‘squeezed’ during the movement process, which was referred to as ‘spring’ characteristics in this research. Further analysis indicated that the characteristics of shock train movement were closely related to the background flow structure, or rather the surface pressure gradient. The shock train leading edges jumped forward in an adverse wall pressure gradient and crept forward in a favorable wall pressure gradient. The whole movement process consisted of 5 motion periods. In the 2nd motion period, speed of leading edges and wave-front Mach number varied, leading to the periodical changes of shock train structures. In the theoretical analysis, shock polar method was applied to clarify the intensity variation of reflection shocks in the shock train.

Direct numerical simulation of flame stabilization assisted by autoignition in a reheat gas turbine combustor

A three-dimensional direct numerical simulation (DNS) is performed for a turbulent hydrogen-air flame, represented with detailed chemistry, stabilized in a model gas-turbine combustor. The combustor geometry consists of a mixing duct followed by a sudden expansion and a combustion chamber, which represents a geometrically simplified version of Ansaldo Energia’s GT26/GT36 sequential combustor design. In this configuration, a very lean blend of hydrogen and vitiated air is prepared in the mixing duct and convected into the combustion chamber, where the residence time from the inlet of the mixing duct to the combustion chamber is designed to coincide with the ignition delay time of the mixture. The results show that when the flame is stabilized at its design position, combustion occurs due to both autoignition and flame propagation (deflagration) modes at different locations within the combustion chamber. A chemical explosive mode analysis (CEMA) reveals that most of the fuel is consumed due to autoignition in the bulk-flow along the centerline of the combustor, and lower amounts of fuel are consumed by flame propagation near the corners of the sudden expansion, where the unburnt temperature is reduced by the thermal wall boundary layers. An unstable operating condition is also identified, wherein periodic auto-ignition events occur within the mixing duct. These events appear upstream of the intended stabilization position, due to positive temperature fluctuations induced by pressure waves originating from within the combustion chamber. The present DNS investigation represents the initial step of a comprehensive research effort aimed at gaining detailed physical insight into the rate-limiting processes that govern the sequential combustor behavior and avoid the insurgence of the off-design auto-ignition events.

Gas-solid flow through the mixing duct and tail section of ejectors: Experimental studies

Detailed experimental investigation has been conducted to study the single-phase (air-air) and two-phase (air-solid) flows through ejectors. An attempt is made to increase the amount of transported solids by introducing design modifications to the mixing section of the ejector. The mixing duct has been augmented by a tail section. Three different geometries of the mixing duct with tail section of ejector have been designed, fabricated and tested experimentally. The effects of the mixing duct and tail section shapes on the ejector performance have been investigated. In addition, the effects of air motive pressure and the solid particles mass flow rate on the static pressure distribution and the vacuum pressure produced have been also studied. The obtained results show that the geometry of convergent-constant-divergent for the mixing duct gives higher vacuum pressure and favourable performance of the ejector.

Experimental investigation of flow behavior for scalloped and lobed mixers

Particle image velocimetry and planar laser-induced fluorescence techniques are used to measure the flow of scalloped, lobed, and splitter mixers. Special attention is given to the understanding of the details of the flow for the scalloped mixers and the role of the notches in enhancing the mixing process. The comparisons between the three mixers are explained by providing detailed measurements on several axial and cross-streamwise planes. Scalloped–lobed mixers enhance the mixing of the two streams by generating an additional pair of small-scale streamwise vortices not observed in the lobed mixer shear layer. Under the same test condition, the impingement of the inner flow on the mixing duct occurs further downstream than the lobed mixer. For the scalloped mixer, the decay of the cross-streamwise velocity starts at its exit plane, while for the lobed mixer, there is an initial magnitude rise before the streamwise decay.