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Fuel injection location studies on pylon-cavity aided jet in supersonic crossflow
Abstract
The current study numerically investigates the effect of fuel injection locations within a pylon-cavity aided Supersonic Combustion Ramjet (SCRAMJET) combustor on mixing enhancement, flame holding, fuel jet penetration and total pressure loss. RANS equations for compressed real gas are solved by coupled, implicit, second-order upwind solver. Two-equation SST model is used for turbulence modelling. The computational model is validated using experimental steady wall pressure data and 2D velocity field. The study uses seven distinct sonic fuel injection location cases of hydrogen fuel through a 1 mm diameter hole along the axis of the test section floor. All cases maintain crossflow of Mach number 2.2. The simulations show that the counter rotating vortex pair within the cavity plays a vital role in fuel dispersion and fuel jet penetration capability. The presence of pylon resulted in an increase of pressure loss by 7%, whereas the influence on total pressure loss due to transverse fuel injection is found to be insignificant. The injection locations within the cavity give around 55% (max) increase in fuel dispersion compared to location upstream of the pylon. Also the cavity floor locations give about 55% - 90% more flammable plume area than the injection from other locations.
... It solves the transport equations for the turbulent kinetic energy κ and for the turbulence inverse time-scale, ω. Turbulence kinetic energy transport equation is, Figure 7. Boundary conditions for the computational domain (28) . and turbulence inverse time-scale transport equation is, ...
... Figure 7 shows the boundary conditions applied for the computational domain. Tables. 2 and 3 give the details regarding the boundary conditions and grid resolution, respectively (28) . Figure 8 shows the grid independence study on local and global variables over five different meshes with varying grid density and wall spacing. ...
... · · · (9) Figure 8. (a) Local pressure variation along the centre line of the test section (b) mixing efficiency (c) mean X-velocity d) H 2 mass fraction variations for various grid refinements for the baseline geometry case (28) . ...
Numerical investigation of the effect of pylon geometry within a pylon-cavity aided Supersonic Combustion Ramjet (SCRAMJET) combustor on mixing enhancement, flame-holding capability, fuel jet penetration and total pressure loss are conducted in the current study. RANS equations for compressed real gas are solved by coupled, implicit, second-order upwind solver. A two-equation SST model is used for turbulence modelling. Validation of the computational model is performed with the help of experimental data collected using surface pressure taps, Schlieren flow visualisation and particle image velocimetry (PIV). The study uses four distinct pylon geometry cases, which include the baseline geometry. Sonic injection of hydrogen fuel through a 1mm diameter hole at 2mm downstream of the pylon rear face along the axis of the test section floor is performed for every case. A crossflow of Mach number 2.2 at four bar absolute pressure and standard atmospheric temperature is maintained. A comparative study of mixing efficiency, total pressure loss, fuel jet penetration and fuel plume area fraction for the different cases evaluate the mixing performance. The simulations show that the Pylon 2 case gives a significant improvement in the performance parameters compared to the other geometries. It is observed that mixing efficiency and fuel jet penetration capability of the system are highly dependent on the streamwise vortex within the flameholder.
... Various shapes of cavity flame holders and their influence on mixing efficiency in a scramjet engine were investigated by Moradi et al. 22 Various studies have also investigated the effect of incorporating a pylon upstream to the cavity injection. [23][24][25] It was found that the fluid mass exchange to the cavity significantly increased (three times) with an upstream ramp when compared to a cavity without a pylon. 26 Oamjee et al. 23 investigated various injection locations inside cavity-pylon setup. ...
... [23][24][25] It was found that the fluid mass exchange to the cavity significantly increased (three times) with an upstream ramp when compared to a cavity without a pylon. 26 Oamjee et al. 23 investigated various injection locations inside cavity-pylon setup. They found that the fuel injection into the cavity resulted in a 55% increase in fuel dispersion compared to fuel injection from the upstream side of the pylon. ...
This study investigates the fluid dynamics and mixing characteristics of an oscillating sonic jet injected into a supersonic cross flow of Mach 2.1 using experimental and computational techniques. The oscillating jet is produced by a novel fluidic oscillator, which consists of a primary rectangular duct that expands into an outer duct with sudden expansion. Control jets are injected in the lateral direction from the side walls of the sudden expansion in an out-of-phase manner to oscillate the injected jet in the spanwise direction of the crossflow. Experimental and numerical investigations based on wall static pressure and mass fraction fluctuations, respectively, revealed that the injected jet oscillation frequency matches the control jet frequency. The iso-surface of lambda-2 criterion showed the presence of various dominant vortex structures, such as counter-rotating vortex pairs, horseshoe vortex, sidewall vortices, and trailing vortices. Helicity contour plots showed that the streamwise vortices oscillate in the spanwise direction with the control strategy and promote the spread of the injected jet in the spanwise direction. The spatiotemporal reconstruction (z–t plot) of the density gradients at a particular streamwise location revealed that the bow shock produced by the interaction of the injected jet and the crossflow oscillates with the actuation of the control strategy. The power spectral density of the z–t plot revealed that the shock wave oscillation frequency matches the control jet frequency. The oscillating jet produced by the control strategy showed significant mixing enhancement in supersonic crossflow compared to a simple rectangular injection.
... A pyramidal-shaped pylon that has 45⁰ leading edge angle with height, length, and width (maximum) of 10 mm, 10 mm, and 5 mm, respectively, is located at the leading edge of the cavity. The earlier studies have predicted the locations C and E which are at a distance of x/d = 25 and 34, respectively, as the best location for fuel injection based on the evaluation of the performance parameters (Oamjee and Sadanandan 2019). So, at locations C and E, cold flow studies are conducted with air injection at sonic conditions and hydrogen fuel injection at reactive flow conditions with a stagnation temperature of 1772.1 K. ...
... The most critical part of a PIV measurement is the seeding particle selection. The flow is seeded with olive oil and the average particle diameter measured using Phase Doppler Particle Analyzer (PDPA) has confirmed that about 85% of the particles generated are about 0.7 µm in diameter (Oamjee and Sadanandan 2019). The region of interest (ROI) for the current study is selected according to the Stokes number calculated. ...
A comparative study is performed to investigate the effectiveness of non-reactive supersonic flow simulations in qualitatively predicting the mixing performance and flameholding capability of a pylon-cavity aided, sonic H 2 fuel injection flameholder under, reactive flow conditions. The performance parameters such as mixing and combustion efficiencies, and flammable plume area are taken into account for the study. A non-reactive, steady-state RANS simulation is solved using coupled, implicit, second-order upwind solver with a two-equation SST κ-ω turbulence model. The numerical scheme is validated experimentally using steady wall pressure data and 2-D velocity vector field. Similar steady-state modeling has been performed for a reactive flow simulation with an 18 step Jachimowski reaction scheme for H 2 -air reactants. The study consists of two distinct injection locations on the cavity floor. Inlet Mach number of 2.2 is maintained for all the cases. The study shows that the cavity vortex pair with recirculating hot burned gases plays a decisive role in accurately predicting the flame location within the configuration. Though the non-reactive flow simulations are helpful in understanding the fundamental mechanisms also relevant under reactive flow conditions, like fuel-air mixing for example, it fails in the accurate prediction of the ignition location and flame stabilization.
... Attaining steady combustion during hypersonic speeds is a challenge due to the rapid air velocity. To address this, various geometric shapes, including cavities [8][9][10][11][12], struts [13][14][15][16][17][18], pylons [19][20][21][22], and shockwave generators [23,24], have been employed within the combustor. These design modifications aim to enhance mixing and combustion processes, mit- ...
The present study explores the characteristics of reacting flow in a scramjet combustor with struts, focusing particularly on implementing different injection strategies. A three-dimensional DLR scramjet combustor is utilised to assess the impact on the system, incorporating multiple injections and varying injection angles on the triangular wedge. The analysis considers three injectors with parallel, upward and downward injections at angles of 15° and 30°. The numerical investigation is con-ducted under a constant total pressure of 7.82 bar, a temperature of 340 K, and an airspeed of Mach 2 at the inlet. The results highlight the significance of injector location and shape in promoting flame stabilization. Furthermore, injection angles play a crucial role in mitigating shockwave intensity. The numerical analysis involves a steady-state Reynolds-averaged Navier-Stokes equation with the shear stress transport k–ω turbulence model. The obtained results were analyzed by examining the critical variables such as Mach number, static pressure and combustion efficiency across the combustor. Based on the com-putational results, injecting fuel upward not only increases the overall pressure loss but also enhances the subsonic regime downstream of the strut, which leads to better mixing and combustion efficiencies. This is primarily due to shockwave generation from the edges of the strut and the interactions with the fuel stream shear layers.
... (Huang et al. 2011;Fuller et al. 1998;Gruber et al. 2004). Compared with wall injection, these devices can effectively improve the fuel mixing efficiency, but their geometry structure hinders the air flow in the combustion chamber, resulting in greater total pressure loss and higher requirements for structural thermal protection (Aryadutt et al. 2019). In addition, related studies have shown that the wall structure of the isolation section has a great influence on the degree of the fuel mixing. ...
Metal powder fuel has high energy and volume calorific values, and it is an excellent power fuel for hypersonic aircraft in the future. However, metal powder has poor flow characteristics and is difficult to be effectively mixed with air when supersonic air is injected. The mixing degree of fuel and air is an important factor to the efficiency of the engine. One way to solve this problem is to introduce a jet (the primary jet) upstream of the powder fuel inlet to improve the mixing degree of fuel and increase the residence time. In this study, a double-hole jet two-dimensional scramjet combustor structure was established, the effects of different initial jet conditions on the performance of scramjets was studied and analyzed. The results show that the introduction of the primary jet can effectively improve the mixing degree of the powdered fuel and air in the combustion chamber, prolong the residence time of the particles in the combustion chamber and improves the combustion efficiency of the metal powder. However, both the combustion efficiency of the metal powder fuel and engine thrust increase and then decrease when the mass flow rate of the primary jet increases.
... The rectangular cavity results in very strong compression corners. The effect of the length to depth ratio (L/D) studied [4] and suggested that the open cavity with the angled aft wall is suitable such that there will be no shocks within the cavity and the circular cavity itself acts as the aft wall to prevent shocks inside the cavity [5]. A closed cavity tends to move the shear stress layer below the 40% of the cavity depth which is not preferable as those conditions tends to produce bow shocks whereas the open cavity maintains the shear stress layer with in the 40% of the depth thereby concluding that the open cavity is suitable for this analysis [6]. ...
Scramjet engines, in the recent times can play a vital role to achieve hypersonic flights. The major drawback is the proper mixture of air and fuel in the combustion chamber. Cavity based mixing have been a promising effect to achieve proper air fuel mixture. This paper discusses 2D and the 3D computational fluid analysis carried out over the various geometries of pylon based trapezoidal cavities using both the methods of k-omega. Second order upwind, coupled, and implicit solver were used to solve the density-based flow. Turbulence modelling were defined by SST model. The results obtained for 2.5 Mach were pressure loss, turbulence intensity and velocity vectors are compared to investigate to find the geometry providing efficient mixing. The tall pylon with a cut geometry among five proposed geometries have shown the efficient results for proper mixing.
... It has achieved reasonable results with RANS method in the similar research, e.g. cantilever ramps [46], pylon-cavities [47], and divergent ramps [48]. ...
The circulation growth of supersonic streamwise vortex (SSV) formed by a ramp is complicated and essential to estimate scramjet mixing behavior, while a predictive model for circulation growth is still lacking. Through well-validated RANS simulations, it is observed that the circulation growth of streamwise vortex for a ramp is similar to one for vortex generator, which determines three important geometric parameters for SSV growth, i.e., ramp angle, swept angle, and length of the leading-edge. The results of different geometries with these three parameters further show that the SSV circulation growth is proportional to the downwash speed influenced by oblique shock intensity. By building the relation between downwash speed and geometric parameters, an SSV circulation prediction model is proposed and validated by the circulation curves of all cases under a wide range of free stream Mach numbers. The novel circulation model has the potential for the preliminary mixing enhancement design for a scramjet ramp injector.
... Therefore, it is necessary to model and calculate the flow field to obtain the corresponding density field distribution. The Reynolds-averaged Navier-Stokes (RANS) solver was used to calculate the steady-state flow-field distribution [19,20]. The head shape of the aircraft is shown in Fig. 3(a). ...
When an aircraft loaded with pulsed laser radar flies at supersonic speed, the laser beam will be distorted by the uneven outflow field, resulting in a significant reduction in ranging accuracy. In this study, the influence mechanism of the shock wave on the performance of forward pulsed laser radar is investigated. First, a novel semi-analytical method is proposed to model the pulsed laser echo wave affected by shock waves, which combines the laser radar equation with optical distortion parameters. Second, an improved ray tracing method based on inverse distance-weighted interpolation with a quadrilateral mesh is proposed to trace the trajectory of the laser beam passing through the flow field, and the effectiveness and superiority of the algorithm are verified. Thereafter, an evaluation method based on the optimal confidence interval is proposed to evaluate the ranging error of pulsed laser radar; which can effectively evaluate the ranging accuracy of pulsed laser radar under the influence of the shock wave. The simulation results show that the ranging performance of pulsed laser radar below Mach 3 is slightly affected, and the detection system error and random error reach the minimum and maximum at Mach 4, respectively. This study provides a theoretical basis for the suppression of the aero-optical effect of forward pulsed laser radar at supersonic speed.
... Works have inclined it at two distinctive angles [21], set it to be 5-10% shallower than the front-wall [22] and also developed it to suppress pressure oscillations by imparting vents, slots and 'beaks' to the rear wall [23]. Pylons have been introduced upstream of the cavity to aid mixing [24,25], while Huang et al. (2013) [26] has performed a multi-objective study of the cavity. However, in all of these studies, no changes to the cavity front-wall geometry were examined. ...
This paper presents experimental data examining the flameholding performance of a scramjet cavity with an inclined front wall. Mach 8 flight-equivalent flows were delivered to the axisymmetric, cavity combustor via the T4 reflected shock tunnel. The combustor model was designed to permit interchangeable cavity geometries. Each examined cavity maintained a length-to-depth ratio of 4, with a depth of 8 mm and cavity close-out angle of 22.5 deg. The first examined cavity possessed a 90 deg front wall/step, while the second examined cavity inclined the front wall at 45 deg. Ethylene fuel was injected at a variety of mass flow rates to examine both scram-mode, jet-wake anchored and dual-mode combustion, with data measurements obtained via wall-mounted static pressure sensors. The steady nature of these combustion modes during the shock tunnel test time was demonstrated. Minimal differences in combustion-induced pressure rises were observed between the two cavity geometries across each fuelling condition. This indicates that the low pressure, quasi-stagnant region downstream of a 90 deg step is not essential to attaining robust flame-anchoring in a scramjet combustor. Chemically reacting, Reynolds-averaged Navier-Stokes computations were performed for the three dimensional combustor. Only minor mixing and combustion performance differences were noted between each geometry/fuelling permutation. It was noted that the 45 deg cavity retained more robust vortex structures within the cavity, with the 90 deg cavity seeing complete breakdown of the cavity vortex structures. This work indicates that the cavity front wall could be further modified with minimal impacts on flameholding performance.
... Li et al. [31] studied the role of the cavity in mixing a liquid jet injected into a supersonic crossflow. Oamjee and Sadanandan [32] numerically investigated the pylon cavity based supersonic combustor for air-fuel mixing. They also studied the location of injection within the cavity to enhance the mixing. ...
Cavities have been extensively studied as flame holding devices for scramjets. Wall cavities are known to have issues of fuel nonuniformity and flame close to the wall. Although strut offers an option to deliver fuel in the core flow, it lacks the mixing efficiency provided by the cavity. This work attempts to combine a strut and a cavity into a strut cavity and compare it with the wall cavity for different Mach numbers and aspect ratios, using numerical simulations. Flow features, vortex movement, and oscillation behavior are investigated and compared. Both cavities exhibited oscillations for lower values of Mach numbers and aspect ratios, while oscillations were found to decrease for high values. The wall cavity showed a higher propensity to oscillations compared to the strut cavity. The process regarding the generation of the waves inside the cavity has been revisited. Both the unstable shear layer and the vortex shedding contribute to the generation of the waves inside the cavity. This merges the previous two theories on the wave generation inside the cavity in the supersonic flows.
... Due to the variance of the fuel penetration capability between the cavity/step and strut/bluff body, the combined physical flameholder can achieve efficient fuel mixing, promote fuel ignition, and extend the flame stability limit. For instance, the cavity-bluff body or cavity-pylon combination flameholding scheme can broaden the operation range of a simple cavity flameholder by instream strut-wake generation [318,319,[378][379][380][381][382][383][384][385][386]. A tandem-type strut-cavity flameholding scheme extends the cavity influence into the core flow [369,387]. ...
A review of fundamental research in combustion stabilization for hypersonic airbreathing propulsion is presented. Combustion in high-speed airbreathing propulsion systems demands stable flame distribution and chemical reaction to provide reliable thrust over a wide flight envelope. Various methods have been developed to stabilize combustion depending on the hypersonic regime. For low hypersonic conditions, combustion occurs mainly in the diffusive mode in which the gas/liquid fuels are injected into supersonic freestreams for simultaneous fuel-air mixing and chemical reaction. Flame stability is generally enhanced by improved mixing, physical flameholding, and external energy addition. In higher hypersonic conditions, partially/fully premixed combustion relying on shock induced stabilization becomes more dominant. In such cases, flame stabilization can be achieved through alternative means such as radical generation and standing oblique detonation waves. The review outlines both experimental and numerical research progress made towards combustion stabilization over the entire hypersonic regime, and intended to lay the groundwork for further studies which can provide optimized design guidelines for the next generation of high-speed airbreathing propulsion systems.
... Alternate approaches to fuel delivery include ramp injectors [2], [3], [4], strut injectors [5], [6], and placement of injectors near acoustic cavities [7], which improves the residence time of fuel injection via the creation of recirculation regions near the injection locations. Additional approaches that have not been studied as extensively include lobed co-axial injection into the supersonic https://doi.org/10.1016/j.ast.2020.105908 ...
The design of scramjet engines includes the investigation of minimally intrusive fuel delivery mechanisms which maximize jet fuel penetration and mixing. In subsonic crossflows it has been shown that pulsation of gaseous jets can improve jet penetration and mixing in comparison to unforced (steady) jets. Extensive investigations have already explored subsonic flow regimes for pulsed injection; however, only a limited number of experimental and numerical studies have been performed which investigate pulsed jets in a supersonic crossflow. This study presents fully three-dimensional (3D) scale-resolving simulations of steady and a sinusoidally pulsed jet (f = 16 kHz) using a wall modeled large eddy simulation approach to capture large-scale vortical structures in turbulent jet in supersonic crossflow and pulsed jet in supersonic crossflow. A block-structured, fully 3D turbulent scale resolving finite volume model was used to explore physics simulations of steady and pulsed jets in a supersonic crossflow. The results of the simulations presented show that sinusoidal pulsation of a gaseous hydrogen jet improves penetration and mixing over the steady jet when the momentum flux ratios are equivalent.
... Transverse fuel injectors are a simple, reliable and conventional method to achieve rapid mixing of air and fuel stream in a SCRAM-JET engine [1][2][3]. Since a SCRAMJET engine operates at very high speeds, the flame-holding and stabilization in such a flow field becomes a very critical issue, which can be solved by placing a backward facing step in the SCRAMJET combustor [4][5][6][7][8][9][10] to act as a flame-holder. The backward facing step creates a large recirculation zone, with the hot gases contained therein, that acts as a continuous ignition zone. ...
A SCRAMJET engine typically has multiple transverse fuel injectors with a flame holder. In this study, we consider a two-jet SCRAMJET engine design configuration that uses the low speed recirculation region created by a backward facing step as a flame holder. The effect of spacing between the transverse fuel injectors on the performance of a SCRAMJET engine has been studied using cold-flow simulations. The position of the leading jet is kept fixed at the end of the recirculation zone as suggested by previous studies and the second jet is placed at various locations downstream in the distinct flow regions formed behind the leading jet. It is assumed that the two jets are identical in dimensions and flow. The spacing between the jets is expected to play a significant role in determine the performance of the SCRAMJET engine. Three-dimensional simulations have been performed, using Menter's SST model in our in-house parallel 3-D RANS unstructured grid CFD solver. The mixing of inlet air and the injected air-fuel in such a SCRAMJET configuration is augmented by the interaction of the transverse under-expanded jet with the incoming supersonic cross-flow through the generation of strong streamwise vorticity. The performance and mixing of the combustor have been quantified for each of the distinct configurations. It is observed that they are indeed affected by the spacing between the jets. From the results presented in this paper, the optimal location for the second jet is at the end of the zone over which the lateral momentum of the first jet is dominant in affecting the jet penetration into the streamwise flow.
... A number of studies examined more novel changes to the cavity geometry within supersonic flows. Cavities zig-zagging across a cavity span [15], 'swallow-tail' cavities [16], pylons introduced upstream of the cavity [17,18], inclining the cavity aft wall at two successive angles [19] and cavities in which the downstream-wall is 5-10% shallower than the upstream-wall [20] have also been examined. Each study varied parameters such as the length-todepth ratios, the downstream-wall angle and the downstream-wall height on the flow within the cavity, while Huang et al. (2013) [21] performed a multi-objective study of the cavity. ...
The cavity flameholder has been widely used in scramjets to improve mixing and combustion rates. While many have investigated modifications to the cavity downstream-wall, few have investigated modifying the upstream-wall - a rear-facing step. The rapid expansion produces a low pressure region and base drag force, while only the larger of the induced twin vortices contributes substantially to mixing. In this paper, the upstream-wall of a scramjet combustor cavity was inclined, aiming to reduce cavity base drag and remove superfluous cavity vortex structures. Upstream-wall angles of 90°, 45° and 22.5° were examined, at cavity length-to-depth (L/D) ratios from 4 to 7. Cavity downstream-wall angles of 90° and 22.5° were examined for each configuration. Reflected shock tunnel experiments were conducted at Mach 7 enthalpy, scramjet combustor conditions. Experimental Schlieren imaging showed that reducing the upstream-wall angle did not significantly affect shear layer separation for L/D ratios of 4 and 5. At L/D ratios 6 and 7, however, reducing the upstream-wall angle saw the shear layer penetrating deeper into the cavity. Reynolds-averaged Navier-Stokes computations examined the internal flow structure and it was shown that reducing the upstream-wall angle to 45° retained the dominant mixing vortex in the cavity, with base drag reduced by 21% compared to the standard 90° upstream-wall case.
... When using CFD method to calculate the density contour, there are three common solvers including Direct Numerical Simulation (DNS), Large Eddy Simulation (LES) and Reynolds Averaged Navier-Stokes (RANS) equations [21][22][23][24]. Based on the instantaneous governing equations, both DNS and LES provide the temporal flow characteristics, which are computationally expensive when compared to RANS. ...
... 'Swallow-tail' cavities displayed improved mass exchange, driven by the 3D vortex ring generated [10]. Pylons introduced upstream of the cavity increased fuel penetration with minimal shock losses [11,12], and improved the exchange of cavity fluid with the freestream [13]. Inclining the cavity aft wall at two successive angles in Mach 1.3 flow reduced pressure losses with reductions in the secondary aft ramp angle [14]. ...
This paper numerically examines methods to modify the upstream wall of cavities within scramjet combustors. The prevailing geometrical features of cavities have remained unchanged since 1957, when the baseline cavity operated in subsonic flows. The rapid expansion experienced over the front wall (a rear-facing step) produces a low-pressure region and accompanying base drag, while the two-vortex structure has only the larger contributing to mixing. The current work examines four new cavity profiles, substituting the rear-facing step for ramps (inclined at 45 and 22.5 deg), or streamtracing it to aerodynamically conform with the primary vortex. Examined in Mach 8 combustor flows, unfuelled flow fields see complete removal of the secondary vortex. A 45 deg ramp angle ensured immediate boundary layer separation. While all cases exhibited reduced drag due to increased pressure loads upon the front wall, combining streamtracing with the ramped profiles provided drag reductions of 12%. Ethylene-fuelled cases exhibited drag reductions to 6.6% for the 45 deg ramp. The stronger single-vortex structure displayed universally improved mixing, with improvements to 16.8%. The 22.5 deg ramp suffered reduced mixing, owing to the shortened residence time experienced. Hence, a minimum front wall angle likely exists. Within lower Mach number flows, universal drag reductions for the unfuelled and fuelled cases were realised, while mixing was comparable amongst most cases. Aerodynamic contouring may hence become more necessary as flight speeds reach Mach 8.
This study explores the flow physics and mixing characteristics produced by the interaction of two transversally injected jets along the spanwise direction with a supersonic crossflow using experimental and numerical techniques. Results show that smaller interjet spacing reduces the crossflow entrainment from the top into the interjet region, while wider spacing facilitates greater entrainment. The investigation further reveals that the gap created by the two injected jets allows the crossflow to decelerate and then accelerate to a higher supersonic Mach number as it flows along the interjet region. Investigation of the shock structures revealed that in the case of smaller interjet spacing, the strong part of the two bow shocks created by the transverse jets interacts with each other, whereas with large interjet spacing, the weak part of the two bow shock waves interacts. This leads to a larger local pressure jump in the interjet spacing with a smaller injector gap than the larger one. Various streamwise vortices, such as horseshoe and counter-rotating vortex pairs (CVPs), are seen to form for spanwise tandem injection in crossflow. The interaction of such vortices is seen to be significant in the case with smaller injector spacing compared to the larger one. The oil flow visualization reveals the formation of a herringbone-shaped separation region in the wake of the jets, and the size of this separation zone diminishes with the reduction in injector spacing. The mixing characteristics investigated using Mie scattering and computations reveal that with an increase in injector spacing, the mixing efficiency increases.
The cavity-assisted scramjet has been proven to be the most promising propulsion system for air-breathing hypersonic vehicles. In this paper, numerical simulations of a Mach 8 axisymmetric scramjet combustor are conducted and validated to investigate the effect of the cavity. The results indicate that the combustion state undergoes significant changes as the combustion heat release increases. Detailed analysis reveals that the role of the cavity in flame stabilization and combustion enhancement also changes with combustion heat release. Under weak heat release conditions, the high-speed environment results in reduced combustion efficiency, and the primary role of the cavity is to stabilize the flame. Increasing the cavity size does not yield significant gains but could bring redundant mass. As heat release intensifies, the combustion enhancement effect of the cavity becomes more prominent. The presence of the cavity dramatically improves fuel combustion efficiency. The distribution of supersonic and subsonic combustion modes, as well as that of premixed and diffusion combustion modes, is also affected by cavity size and combustion heat release. In the engineering development of scramjets, it is suggested that the design of the cavity flameholder should involve careful consideration of combustion heat release.
Owing to the safety concerns associated with the usage of hydrogen (H2) fuel in conducting experiments, the current study investigates the suitability of using helium (He) gas as a surrogate fuel for H2 for supersonic mixing studies. The study adopts a steady state computational approach to compare the mixing performance between the two injection cases, where a wall-based pylon-cavity aided flameholder configuration is used for the investigation. A common fuel injection location at the cavity floor is used for He and H2 injection cases with an injector hole diameter of 1 mm. An inflow Mach number of 2.2 with sonic transverse fuel injection is kept the same for all the cases. A steady Reynolds-averaged Navier–Stokes (RANS) based computational model for compressed real gas is used for the investigation. The governing equations are solved by an implicit, second-order upwind solver with a two-equation Menter's Shear Stress Transport (SST) turbulence model. The study shows a variation in near field mixing performance between He and H2 injection cases due to the difference in the molecular mass. As a result, a 10%–11% variation in the mixing efficiency is observed between the two cases. This makes He not suitable for micro-level mixing studies as a surrogate to H2. However, it is possible to closely predict the trend in global mixing performance parameters as in the H2 injection case.
One of the concerns in the process of measuring laser radar cross-section (LRCS) using optical measurement methods is the aero-optical effect around a high-speed flow field. In this paper, a computational model for LRCS considering supersonic flow in a non-homogeneous medium is presented. First, by taking a typical 15° half-cone blunt cone as the research object, based on CFD calculation, the flow field distribution of the non-uniform medium around the high-speed vehicle is obtained. The ray-tracing method is used to calculate the light transmission in an aero-optical environment. An LRCS approach for arbitrarily complex targets in the aero-optical environment is then proposed, and the method is verified by analytical and numerical methods. Finally, the effects of incoming density, incoming Mach number, and material bidirectional reflectance distribution function (BRDF) parameters on LRCS measurements are investigated, and the errors caused by the steady aero-optical environment are analyzed. The results show that the maximum relative error of the measurement caused by the aero-optical environment in the direction of the shock-wave layer observation is significant when the Fresnel function approximation exponential function parameter is too large or when the root mean square (RMS) of the surface slope is too small in the BRDF parameters of the material. The influence of the aero-optical environmental flow field on the LRCS is characterized by a directional and regional distribution, determined by the incoming Mach number and shock-wave angle, with the influence of the incident flow density on the regional distribution being weakly correlated. The proposed method and calculation results are useful for the simulation, analysis, and calibration of LRCS measurements in complex aero-optical environments.
In this paper, the combustion characteristics of kerosene-fueled supersonic combustor under the conditions of Mach number 2.0, the total temperature at 700 K and the total pressure at 520 kPa (simulated flight Mach number at 3.5) were studied by using the flame stabilizing method of cavity and strut from three aspects such as blockage ratios, kerosene equivalence ratios, location and quantity of injection holes. The results showed that: (A) The combustor with the strut realized the independent and stable combustion of kerosene. The combustion-induced back pressure in the block test with blockage ratio of 20% and 10% destroyed the inlet flow conditions; while the blockage ratio was 7.3% and 5%, the incoming flow conditions when kerosene was burned stably were not destroyed. (B) The kerosene Equivalence Ratio (ER) was more likely to be disturbed upstream than the induced back pressure when it rose, and an excessively high-ER would reduce the combustion efficiency; when the equivalence ratio was constant, the combustion efficiency of the blockage ratio of 7.3% was higher than the blockage ratio of 5%. The combustion efficiencies were 0.86 (ER = 0.13) and 0.78 (ER = 0.19) when the blockage ratio was 5%, respectively; the combustion efficiencies were 0.89 (ER = 0.16) and 0.82 (ER = 0.19) when the blockage ratio was 7.3%, respectively; the combustion efficiencies were 0.51 (ER = 0.25), 0.81 (ER = 0.3), 0.65 (ER = 0.34) and 0.62 (ER = 0.42) when the blockage ratio was 20%, respectively. (C) The porous injection provided behind the strut was beneficial to the atomization of kerosene and improved the combustion efficiency of kerosene; the second injection after the cavity would reduce the combustion efficiency due to insufficient oxygen and combustion space. This study expanded the working range of ramjet and provided a reference fuel injection scheme for Turbine-Based Combined Cycle (TBCC) engine.
Supersonic turbulent non-reacting flow through a cavity-type flameholder employed in scramjet engines with and without subcavities is numerically investigated. The governing equations describing the flow are solved by employing a density-based solver i.e., rhoCentralFoam which is an open-source computational fluid dynamics (CFD) code in OpenFOAM. The turbulence is modeled using the two-equation k-ω SST (shear stress transport) turbulence model. In the present study implication of subcavity types (i.e., extrusive and intrusive) on recirculation patterns and their strength at different subcavity aspect ratios (l/d) and subcavity locations at different inflow Mach numbers is investigated. A significant increase in the strength of the primary recirculation zone is observed for modified rectangle cavity 2 (MRC2) compared to the base rectangle cavity (BRC) and modified rectangle cavity 1 (MRC1). Results show a significant decrease in the size of the secondary recirculation zone for modified angle cavity 2 (MAC2) for all the subcavity aspect ratios and subcavity leading-edge distances (SLDs) at Mach 2. Results also indicate that at a fixed Mach number, the strength of the primary recirculation decreases with the increase in base cavity aspect ratio (L/D) for all the cavity types. Results show a substantial decrease in the velocity magnitude of cavity fluid with the decrease in cavity aft wall angle from 90° to 30°, which shows that the strength of the primary recirculation zone for BRC and MRC2 is higher compared to base angle cavity (BAC) and MAC2. Results show a 20.4776 % and 7.3822 % increase in the peak values of streamwise and normal velocities for MRC2 compared to BRC and MRC1 at Mach 2, and the corresponding increase in velocities for Mach 3 are 14.7979% and 5.6693% respectively for all the aspect rations and SLDs. Present results are validated with the experimental results available in the literature.
A strut/wall combined fuel injection scheme was experimentally investigated to improve the combustion performance in a flush-wall scramjet combustor fueled with liquid kerosene. A thin strut was adopted as fuel injector and flame holder, and wall injectors were also placed on side wall of the combustor, downstream of the strut. With the injection of pure oxygen through the injectors in the trailing edge of the strut, a stable flame was achieved in the combustor, and the flame interaction characteristics were experimentally investigated. Experiments were conducted at the inlet Mach number of 2.0 with different strut/wall fuel distribution schemes. Flame was divided into two parts, namely the core flame and the wall flame, and interaction mechanism between the two parts was experimentally investigated. Experiment results show that core flame is enhanced with the increasing of fuel equivalence ratio. Core flame propagates to side wall of the combustor to ignite wall fuel, forming a wall flame. Wall flame leads to a thermal chocking downstream of strut to enhance core flame. Combustor performance and flame characteristics are influenced by fuel distribution schemes. The location of wall fuel injectors is also optimized to improve combustion efficiency in this paper. These results in this paper are valuable for the future optimization of combustion organization in scramjet combustor with multi-stages fuel injection scheme.
Experiments are performed to investigate the near blowoff characteristics of cavity-stabilized flames in an ethylene-fueled scramjet combustor. Heated air enters the combustor with conditions of Ma = 2.52, stagnation pressure p=01.6 Mpa and stagnation temperature T0= 1400 K. The lean blowoff and ignition limits are obtained for different-sized cavity flameholders with upstream flush-wall fuel injection, and the flame structures and dynamics near blowoff are captured with high-speed chemiluminescence imaging. Though the flow residence time is believed to increase with increasing cavity size, it is observed that the lean ignition and blowoff limits do not vary monotonically with the cavity size. Under the same fuel injection conditions, the flame intensity in larger cavities is weaker than that in smaller cavities though the residual flame can survive for an obviously longer period in larger cavities. This may be attributed to the leaner mixture within larger cavities when the lean blowoff limits are approached. Thus, there exists an optimal cavity size which is the most suitable for the ignition and flame stabilization near lean blowoff conditions. This optimal cavity size is determined by the competition between the flow residence time and the local equivalence ratio that greatly depends on the interactions between the fuel jet and the cavity.
Base on the concept of molecular mean free paths in high-altitude rarefied aerodynamics, a novel conceptual lightweight design is proposed and applied to the aerodynamic design of wings of hypersonic vehicles cruising in the upper atmosphere between 120 km to 300 km. The lightweight scheme proposed here is promising in hypersonic vehicles cruising in the upper atmosphere in the sense that it is able to not only reduce the weight of a wing by about 14%, but also save the material by the same amount. A comprehensive aerodynamic analysis indicates that the lightweight aerodynamic design can apply to a wide range of cruising altitudes and Mach numbers. In comparison with the conventional flat-plate wing, the lightweight wing always has smaller drags and larger effective lift-to-drag ratios, proving the feasibility of lightweight design for hypersonic vehicles at high altitudes. Moreover, as the cruising altitude is increased, the lightweight wing has even better lift and lift-to-drag-ratio performances, further demonstrating the excellent promise of lightweight wings for hypersonic vehicles cruising in the upper atmosphere. This lightweight design concept is of great value because the region of the upper atmosphere is relatively unexplored, and hypersonic vehicles cruising in this region can be used for high-resolution earth surveillance, accurate gravity/magnetic field mapping, and global ocean climate exploring.
To facilitate burning of ethylene fuel in a scramjet combustor, as a relevant route to hydrocarbon combustion, we propose an alternative design of configuration. Three-dimensional flow field in a strut-based scramjet combustor has been investigated computationally. The Large Eddy Simulation (LES) turbulence model and Partially Stirred Reactor (PaSR) combustion model are used. Validations have been performed against published literature to ensure the numerical settings and accuracy. It is shown that, while hydrogen can be successfully ignited and lead to stabilized flame, ethylene combustion cannot be sustained in the same way. By adding a porous cylindrical burner at downstream of the strut injector, consequently, continuous ignition and flame sustainment of ethylene in the primitive combustor can be achieved. To understand the essential difference, the key factors such as net thrust, momentum and pressure variations, flame extinction, and flame stabilization mechanisms have been studied. Specifically, while the porous cylinder may provide a re-ignition source for the hot flow quenched upstream, it expectably introduces a bow shock and extra drag when being placed at the downstream supersonic region. When the porous burner is inserted right behind the strut where the flow is locally subsonic, however, no shock is formed additionally and the total drag is not significantly increased while flames can be stabilized downstream of the porous burner. As a consequence, a positive net thrust and specific impulse at the same flow rate of fuel mass can be obtained, in contrast to the decelerating state in the primitive configuration that yields a negative thrust. Furthermore, the auxiliary device renders a larger penetrating depth of the injected fuel and flame spreading area as well as higher consumption ratio of oxygen, leading to a more favorable condition of design for a supersonic nozzle to produce effective propulsion.
The effect of expansion fan on cavity flameholding was investigated in a rocket-based combined cycle combustor fueled by ethylene. The inflow of the combustor was at a condition where the Mach number, stagnation temperature, and total pressure were 2.92, 1650 K, and 2.6 MPa, respectively. The global equivalence ratio was 0.25. A backward-facing step (BFS) was used to simulate a rocket that was shut down and create an expansion fan. The dimensionless distance between the BFS and the cavity (d/H) varied from 1.5 to 7.5, where d = 60–300 mm was the actual distance and H = 40 mm was the height of the combustor entrance. The result indicated that the cavity shear layer reattached at the cavity floor in the combustor with 2.5 ≤d/H≤ 3.5 because of the pressure gradient resulting from the expansion fan. The expansion fan also significantly decreased the static pressure in the cavity. As a consequence, the combustor with 2.5 ≤d/H≤ 3.5 failed in ignition. In the combustors with d/H = 4.0 and 6.0, the cavity flows were open-type. However, the ignition attempt failed because the expansion fan and the reattachment shock wave might worsen the local equivalence ratio in the cavity. Successful ignition and flame stabilization were achieved in the combustors with d/H = 1.5, 2.0, 4.5, and 7.5. Three combustion modes and three kinds of combustion oscillation were identified. In the combustor with d/H = 1.5, only a small part of the cavity was exposed to the expansion fan. The combustor operated in the cavity stabilized scramjet mode because the flame front was anchored at the cavity leading edge. The cavity shear layer impingement onto the cavity ramp lead to the combustion oscillation with a frequency ranging from 300 to 400 Hz. The combustion mode and combustion oscillation in the combustor with d/H = 2.0 were the same as their counterparts in the combustor with d/H = 1.5. Unsteady jet-wake stabilized scramjet mode was witnessed in the combustor with d/H = 4.5. In this combustor, the fuel jet and the front of the cavity were exposed to the expansion fan. The reattachment shock wave lifted the cavity shear layer. The combustion oscillation induced by shock-flame interaction showed a dominant frequency at 107.1 Hz. The expansion fan and the reattachment shock wave were upstream of the cavity in the combustor with d/H = 7.5. Both of them made the boundary layer more sensitive to adverse pressure gradient. The flame front was upstream of the cavity because the heat release resulting from vigorous chemical reaction choked the combustor and induced large-scale boundary layer separation. The combustor operated in ramjet mode. The unsteady shock train and thermal throat induced the combustion oscillation with a frequency of approximately 800 Hz.
Since boundary layer combustion is an effective method to reduce skin friction in supersonic flow-fields, effects of the combustion on typical pressure gradient flow fields are numerically investigated using Reynolds-averaged Navier-Stokes (RANS) simulations. RANS method is validated by comparing the numerical results with the conducted experimental data. Wall heat flux of combustion case is reduced rather than increased although the flame is restricted in boundary layer, especially in adverse pressure gradient state. Compared with the results of no-injection cases, great skin-friction reduction can be obtained by boundary layer combustion, and further reduction can be achieved as pressure gradient increase. Then, the reduction mechanism is analyzed in details. Due to low viscosity gas injection and decline of velocity gradient, the skin friction level is reduced. Moreover, the component of shear stress, Reynolds stress, is decreased due to low density caused by combustion. Additionally, wall velocity laws of combustion coupled with pressure gradient are explored. The results show that the corrected White and Christoph's law is appropriate for velocity description even in sub-layer because the parameters used in the law are obtained at wall and modified by combustion heat release. The research popularizes the application of boundary layer combustion to common pressure gradient situation, and gives advice on wall function selection of turbulent boundary layer combustion.
A numerical investigation on the effects of fuel injection angle on various mixing parameters within a pylon-cavity aided supersonic combustor flameholder under non-reactive flow conditions is performed. The computational model based on Reynolds-averaged Navier–Stokes equations for compressed real gas is solved by a coupled, implicit, second-order upwind solver with a two-equation Menter’s shear stress transport turbulence model. The steady simulations are experimentally validated using wall pressure data, two-dimensional (2D) velocity field, and fuel mass fraction. Three distinct fuel injection locations at the cavity floor are used for sonic hydrogen fuel injection at 90° and 45° injection angles, with a crossflow Mach number of 2.2. The results show deeper fuel jet penetration capability for the transverse injection when compared to an angled injection, whereas better mixing capability is observed for the latter. The fuel jet vortex pairs formed due to the interaction of the surrounding cavity flow with the barrel shock play a vital role in the mixing mechanisms. The lower pressure regions due to the barrel shock result in the formation of a secondary fuel jet vortex pair. The Kelvin–Helmholtz instability observed between the counter-rotating vortex pairs results in the formation of smaller eddies, which enhance the fuel dispersion and transport.
In this work, a computational fluid dynamic approach is employed to investigate the flow and mixing characteristics of the fuel jet released from the trailing edge of the lobe strut at the supersonic free stream. Hydrogen is injected at sonic velocity while the supersonic freestream moves parallelly over the strut. The lobe shape of the strut is selected to generate strong streamwise vortices and consequently to boost the air/hydrogen mixing. The size and strength of the vortices are visualized through the 3-D contour of the mixing zone to reveal the chief effective terms on the fuel distribution inside the combustor. RANS equations with the SST turbulence model are used for the computational simulation of the supersonic air-stream over lobe strut. The impacts of free-stream Mach number and pressure of jet on fuel mixing are also examined. Moreover, the vertical and horizontal injection models are compared to reveal the mixing mechanism in the downstream of the lobe strut. Our results indicate that the vertical edge is more effective than the horizontal edge on the efficiency of the mixing zone. According to our findings, raising the free stream velocity increases the strength of the vortices and consequently, fuel mixing enhances in the downstream of the injector.
Penetration and distribution of fuel inside the supersonic combustor significantly influence on the overall performance of the scramjet. This research employed the numerical technique to examine the mixing performance of the multi-hydrogen jets at supersonic airflow when the downstream step exists in the downstream of the jet. CFD simulations are conducted to disclose the feature of multi jets when a downstream step is applied. The effects of back step height, free stream Mach number and fuel jet pressure on the mixing efficiency of the four hydrogen jets are disclosed. Our study indicates that decreasing the back step height from 3 mm to 1.5 mm increases the mixing rate up to 28% in downstream of the multi jets. Besides, the formation of the normal shock in upstream of the jet reduces fuel mixing while the normalized mixing factor of fuel jets enhances in downstream of injectors.
The injection of the fuel is a highly important process for the enhancement of the scramjets. In this article, the presence of the backward-facing step on the mixing of the multi-fuel jets is expansively studied. The primary attention of this article is to scrutinize the flow feature of the fuel jet under the backward-facing step. The mixing mechanism of the fuel is also studied to compare this injection system with conventional methods. To do this, a 3-dimensional model is chosen to consider the real physic of the problem. Reynolds Average Navier-Stocks equations are solved with a computational fluid dynamic method to visualize the flow pattern of the fuel jet at the free stream Mach number of 4. SST turbulence model is also used for the calculation of the viscosity. Our results indicate that increasing the jet space from 4 to 10 times of jet diameter in the presence of the backward-facing step increases the mixing efficiency up to 20% in the downstream. Our findings depict that augmenting the number of fuel injectors from 4 to 8 augments the mixing rate up to 15% inside the combustor.
Due to the wide flammability range and low ignition delay of Hydrogen, it can burn rapidly
and generate a large amount of thrust. It is for this reason alone that many researchers have promoted hydrogen as a potential fuel in scramjet engines. As H 2 is also environmentally safe and clean and can be produced from abundant sources, several researchers across the globesupport the use of H2 fuel. However, its lower volumetric energy density and higher flammability range also have an adverse effect in on-board storage system for aircraft propulsion applications. This review gives a brief representation of Hydrogen fueled scramjet engine as well the challenges associated with H 2 fuel. Additionally, the advantage of hydrogen as a fuel as compared to other hydrocarbon fuel is also discussed here thoroughly.
A comprehensive numerical calculation has been conducted to investigate a rarefied hypersonic flow past a three-dimensional (3-D) cavity with a length-to-depth ratio of 4 using the direct simulation Monte Carlo (DSMC) method. In this calculation, the representative atmospheric environment at altitudes of 70 km, 75 km, 80 km and 90 km is considered and Maxwell model was employed to simulate the gas-surface interaction (GSI). A comprehensive understanding of the effects of Maxwellian accommodation coefficient and free-stream Knudsen number on flow-field characteristics is obtained and the sensitivity of aerodynamic surface quantities i.e. the pressure, shear stress, and heat flux, to the accommodation coefficient and free-stream Knudsen number is assessed in depth based on the flow-field patterns from a point of view of gas kinetic theory. The results emphasise the sensitivity of flow-field characteristics inside the cavity and aerodynamic surface quantities on the cavity surfaces to the accommodation coefficient and free-stream Knudsen number. As the GSI changes from diffuse into specular reflection, the two vortices inside the cavity move closer to each other and merge into one, leading to a representative open cavity. The peak values for pressure on and heat flux to some surfaces of the cavity in the case of near specular reflection are much larger than that in the diffuse-reflection case. As free-stream Knudsen number rises, the ability of external stream to enter the cavity is lowered, causing that the flow pattern inside the cavity develops from a closed flow with two vortices generated into an open flow with only one vortex produced and the peak-density point changes from the lower-right corner to the exit of the cavity. In addition, the increase in free-stream Knudsen number also results in more non-uniform distributions of aerodynamic surface quantities along the aft wall, with most of the pressure and heat loads exerted on the top the aft wall for the case of a largest free-stream Knudsen number considered here.
An efficient mixing process is very important for the engineering implementation of an airbreathing propulsion system. The air and injectant should be mixed sufficiently before entering the combustor. Two new wall-mounted cavity configurations were proposed to enhance the mixing process in a conventional transverse injection flow field. Their flow field properties were compared with those of a system with only transverse injection ports. Grid independency analysis was used to choose a suitable grid scale, and the mixing efficiencies at four cross-sectional planes (namely x=20, 40, 60, and 80 mm, which are just downstream of the jet orifice) were compared for the configurations considered in this study. The results showed that hydrogen penetrated deeper when a cavity was mounted upstream of the transverse injection ports. This is beneficial to the mixing process in supersonic flows. The mixing efficiency of the configuration with the wall-mounted cavity was better than that of the conventional physical model, and the mixing efficiency of the proposed novel physical model I (98.71% at x=20 mm) was the highest of all. In the case with only transverse injection ports, the vortex was broken up by the strong interaction between the shear layer over the cavity and the jet.
Sustained combustion and optimization of combustor are the two challenges being faced by combustion scientists working in the area of supersonic combustion. Thorough mixing, lower stagnation pressure losses, positive thrust and sustained combustion are the key issues in the field of supersonic combustion. Special fluid mechanism is required to achieve good mixing. To induce such mechanisms in supersonic inflows, the fuel injectors should be critically shaped incurring less flow losses. Present investigations are focused on the effect of fuel injection scheme on a model scramjet combustor performance. Ramps at supersonic flow generate axial vortices that help in macro-mixing of fuel with air. Interaction of shocks generated by ramps with the fuel stream generates boro-clinic torque at the air & liquid fuel interface, enhancing micro-mixing. Recirculation zones present in cavities increase the residence time of the combustible mixture. Making use of the advantageous features of both, a ramp-cavity combustor is designed. The combustor has two sections. First, constant height section consists of a backward facing step followed by ramps and cavities on both the top and bottom walls. The ramps are located alternately on top and bottom walls. The complete combustor width is utilized for the cavities. The second section of the combustor is diverging area section. This is provided to avoid thermal choking. In the present work gaseous hydrogen is considered as fuel. This study was mainly focused on the mixing characteristics of four different fuel injection locations. It was found that injecting fuel upstream of the ramp was beneficial from fuel spread point of view.
An investigation of the nonreacting flow associated with pylon-aided gaseous fuel injection into a Mach 2 crossflow is described. In this study, a small pylon was positioned just upstream of a circular flush-wall fuel injector. Three pylon geometries were studied, along with a no-pylon reference case. In all cases, a typical cavity-based flameholder was positioned downstream of the fuel injector. The injectant plume characteristics were interrogated using a variety of laser-based and probe-based measurement techniques. Planar laser-induced fluorescence of nitric oxide was used to study the instantaneous plume structure. Spontaneous vibrational Raman scattering provided time-averaged plume characteristics and mixing information. Probe-based instrumentation was used in conjunction with the mixing data to estimate the total pressure losses associated with each configuration. Each pylon had a unique influence on the fuel-injection plume. In all cases, the presence of the pylon resulted in improved fuel penetration into the supersonic crossflow without significantly changing the total pressure-loss characteristics. Mixing efficiencies of the pylon-aided injection cases were not substantially different from the reference case.
This document describes the current formulation of the SST turbulence models, as well as a number of model enhancements. The model enhancements cover a modified near wall treatment of the equations, which allows for a more flexible grid generation process and a zonal DES formulation, which reduces the problem of grid induced separation for industrial flow simulations. Results for a complete aircraft configuration with and without engine nacelles will be shown.
The present study describes the numerical investigations concerning the combustion enhancement when a cavity is used for the hydrogen fuel injection through a transverse slot nozzle into a supersonic hot air stream. The cavity is of interest because recirculation flow in cavity would provide a stable flame holding while enhancing the rate of mixing or combustion. Several inclined cavities with various aft wall angle, offset ratio and length are evaluated for reactive flow characteristics. The cavity effect is discussed from a viewpoint of total pressure loss and combustion efficiency. The combustor with cavity is found to enhance mixing and combustion while increasing the pressure loss, compared with the case without cavity. But it is noted that there exists an appropriate length of cavity regarding the combustion efficiency and total pressure loss.
Reaction rate coefficients and thermodynamic and transport properties are reviewed and supplemented for the 11-species air model which can be used for analyzing flows in chemical and thermal nonequilibrium up to temperatures of 3000 K. Such flows will likely occur around currently planned and future hypersonic vehicles. Guidelines for determining the state of the surrounding environment are provided. Curve fits are given for the various species properties for their efficient computation in flowfield codes. Approximate and more exact formulas are provided for computing the properties of partially ionized air mixtures in a high energy environment. Limitations of the approximate mixing laws are discussed for a mixture of ionized species. An electron number-density correction for the transport properties of the charged species is obtained. This correction has been generally ignored in the literature.
The computer codes developed provide data to 30000 K for the thermodynamic and transport properties of individual species and reaction rates for the prominent reactions occurring in an 11-species nonequilibrium air model. These properties and the reaction-rate data are computed through the use of curve-fit relations which are functions of temperature (and number density for the equilibrium constant). The curve fits were made using the most accurate data believed available. A detailed review and discussion of the sources and accuracy of the curve-fitted data used herein are given in NASA RP 1232.
In this study, numerical simulations are performed to study the effect of micro air jets on the mixing of fuel in the cavity flameholder of the scramjet. This research mainly focused the optimum position of fuel (C2H4) injection on the mixing rate inside the cavity. In order to simulate the cavity flameholder with micro air/fuel jets, a three-dimensional model is chosen and computational fluid dynamic approach is used for the simulations. The effect of significant parameters is studied by using the Reynolds-averaged Navier–Stokes equations with Menter's Shear Stress Transport (SST) turbulence model. In this work, the transient study is also performed to reveal the flow feature and mass distribution inside the cavity in the supersonic free stream (M=2.2). Results show that the injection of the fuel in the middle of the vertical wall significantly enhances the mixing of fuel in the cavity. The obtained results reveal that the injection of micro air jets distributes the fuel uniformly inside the cavity. Therefore, an enhanced mixing zone occurs in the downstream of the injection slots which lead to flame-holding.
To study the thermal behavior in the cracking reaction zone of regeneratively cooled scramjet cooling channels at different aspect ratios, 3-D model of fuel flow in terms of the fuel's real properties and cracking reaction is built and validated through experiments. The whole cooling channel is divided into non-cracking and cracking reaction zones. Only the cracking reaction zone is studied in this article. The simulation results indicate that the fuel conversion presents a similar distribution with temperature because the fuel conversion in scramjet cooling channels is co-decided by the temperature and velocity but the temperature plays the dominate role. For the cases given in this paper, increasing the channel aspect ratio will increase the pressure drop and it is not beneficial for reducing the wall temperature because of the much severer thermal stratification, larger conversion non-uniformity, the corresponding M-shape velocity profile which will cause local heat transfer deterioration and the decreased chemical heat absorption. And the decreased chemical heat absorption caused by stronger temperature and conversion non-uniformities is bad for the utilization of chemical heat sink, chemical recuperation process and the ignition performance.
The mixing process between the injectant and air is very crucial for the engineering implementation of the scramjet engine, and this is due to the very short resident time of fuel in supersonic flows. In the current study, the three-dimensional Reynolds-average Navier-Stokes (RANS) equations coupled with the two equation . k-ω shear stress transport (SST) turbulence model have been employed to investigate the transverse injection flow field with the pseudo shock wave induced by the high back pressure, and the freestream Mach number is 3.75. At the same time, the influence of the back pressure on the flow field properties has been evaluated as well. The obtained results show that the pseudo shock wave induced by the back pressure plays an important role in the mixing enhancement between the injectant and air. When the back pressure ratio is larger than 5.0, the mixing efficiency increases with the increase of the back pressure ratio. However, when the back pressure ratio is 3.0, the near-field mixing process has been improved, and accordingly its mixing efficiency in this region is larger than the benchmark. This implies that the intense combustion downstream of the injector can enhance the mixing process between the injectant and air, and the mixing and combustion process can be enhanced mutually. When the pseudo shock wave has been pushed upstream of the wall orifice, more injectant has been brought into the separation zone upstream of the injector, and this is beneficial for the mixing process between the injectant and air.
Flows around two-dimensional rectangular cavities driven by thick shear layers are investigated experimentally at two supersonic Mach numbers (M e = 1.5 and 2.5) to show the effects of variations in Mach number and length to depth ratio of the cavity. Flow oscillation is observed in the cavity. The characteristics of the oscillatory behaviour are determined by Mach number and the length of depth ratio of the cavity, as well as the shear layer spanning the cavity. Two oscillatory mechanisms can be identified: one in which a strong trailing-edge vortex and vortices which are shed from the leading-edge interact, the other in which a transverse oscillation of a single vortex occurs within the cavity. Changes in the time-dependent and the time-mean flow characteristics at different flow conditions are discussed.
The time-dependent experimental results are compared with existing theoretical analyses of the frequencies. For one of these characteristic types of oscillation, the longitudinal oscillation, an existing theoretical description is improved with a modified phase relation.
Spark ignition experiments of liquid kerosene are conducted in a scramjet model equipped with dual-cavities at Ma 4.5 flight condition with a stagnation temperature of 1032 K. The ignition ability of two cavities with different length is compared and analyzed based on the wall pressure distribution along the combustor and the thrust evolution. The experimental results indicate that the longer cavity (L/D=7) is more suitable than the smaller cavity (L/D=5) in spark ignition. When employing the smaller cavity, three steady combustion states are observed after spark ignition. The concept of 'local flame' is adopted to explain the expanding problem of weak combustion. The local equivalence ratio in the shear layer is the dominated factor in determining the developing process of local flame. The final steady combustion mode of the combustor is dependent on the flame developing process. When employing the longer cavity, the establishment of intense combustion state can be much easier.
As one of the most promising hypersonic propulsion systems for hypersonic vehicles, the scramjet engine has drawn an ever increasing attention of researchers worldwide. At present, one of the most important issues to be dealt with is how to improve the fuel penetration and mixing efficiency and make the flame stable in supersonic flows. Further, how to reduce the structural weight of the engines is an urgent issue that needs to be considered. The ongoing research efforts on fuel injection techniques in the scramjet engine are described, mainly the cavity flame holder, the backward facing step, the strut injection and the cantilevered ramp injection, and the flow field characteristics and research efforts related to these fuel injection techniques are summarized and compared. Finally, a promising fuel injection technique is discussed, namely a combination of different injection techniques, and the combination of the cantilevered ramp injector and the cavity flame holder is proposed. This is because it can not only stabilize the flame, but also shorten the length of the combustor, thus lighten the weight of the scramjet engines.
An unsteady simulation of a simple axisymmetric inlet-fueled scramjet engine concept is performed using a hybrid Reynolds-averaged Navier-Stokes and large-eddy simulation approach. The freestream has a Mach number of 7.5 with Mach 8 flight enthalpy. The simulation is of a nonreacting case in which hydrogen is injected into nitrogen. The simulation is used to provide a detailed description of the structure of the flow. The simulation shows that a large-scale pair of counter-rotating vortices forms within the scramjet combustor, with rotation opposite to the rotation of the pair that forms further upstream due to the interaction of the fuel plume with the crossflow. This vortex pair is found to significantly alter the shape of the hydrogen fuel plume and increase the rate at which the hydrogen is mixing by more than a factor of 2 compared to before the vortex pair is formed. The distribution of hydrogen is examined in detail. The time-averaged and fluctuating wall pressures, the mean velocity field, and resolved turbulence quantities are also examined. Additionally, the hybrid Reynolds-averaged Navier-Stokes and large-eddy simulation results are used to evaluate the performance of a steady-state Reynolds-averaged Navier-Stokes simulation of the configuration.
This paper presents results from both computational fluid dynamic and wind-tunnel experiments of in-stream fueling pylons injecting air, ethylene, and methane gas into Mach number 2.0 cold airflow. Three fuel-injection pylons studied include a basic pylon, a ramp pylon, and an alternating-wedge pylon. The latter two pylons introduce streamwise vorticity into the flow to increase mixing action. The computational fluid dynamic solution was accomplished using the commercial code FLUENT®. Three wind-tunnel experimental techniques were used: aerothermal probing, Raman spectroscopy, and nitric-oxide planar laser-induced fluorescence. Four measures reported include streamwise vorticity, total-pressure-loss, mixing efficiency, and flammable plume extent. The ramp and alternating-wedge pylons show decisive increases in mixing capability compared with the basic pylon for a finite distance downstream of the injector. The alternating-wedge pylon exhibits a measurable increase in total pressure loss compared with the basic pylon, and the ramp pylon exhibits a negligible increase in total pressure loss compared with the basic pylon. For comparison, the downstream mixing effectiveness of the three pylons is compared with the downstream mixing effectiveness of a transverse circular wall injector studied in past research. In addition, a qualitative comparison between the computational fluid dynamic and wind-tunnel experimental results is made.
This study explores the effect of adding a pylon to the leading edge of a cavity flameholder in a scramjet combustor. Data were obtained through a combination of wind-tunnel experimentation and steady-state computational fluid dynamics. Wind-tunnel data were collected using surface pressure taps, static and total probe data, shadowgraph How visualization, and particle image velocimetry. Computational fluid dynamics models were solved using the commercial FLUENT software. The addition of an intrusive device to the otherwise low-drag cavity flameholder offers a potential means of improving combustor performance by enabling combustion products to propagate into the main combustor flow via the low-pressure region behind the pylon. This study characterized the flowfield effects of adding the pylon as well as the effect of changing Reynolds numbers over the range of approximately 33 x 10(6) to 55 x 10(6) m(-1) at a Mach number of 2. The addition of the pylon resulted in approximately 3 times the mass flow passing through the cavity compared with the cavity with no pylon installed. Reynolds number effects were weak. The addition of the pylon led to the cavity fluid traveling up to the top of the pylon wake and significantly increasing the exposure and exchange of cavity fluid with the main combustor flow.
Fueling the core airflow of a circular or elliptical scramjet combustor cross-section often requires intrusive geometries. Intrusive geometries can distribute the fuel evenly across the combustor cross-section and act as a flameholder for the fuel/air mixture. Compared to conventional transverse or angled wall injection, intrusive geometries allow easier penetration into the core combustor airflow and reduced fuel injection pressures. The design and testing of an intrusive pylon geometry for scramjet combustor fueling is the subject of this research. Three pylon configurations are compared: a basic pylon, a ramp pylon, and an alternating wedge pylon. All three pylon configurations exhibit the same frontal area and inject fuel parallel to the combustor airflow with long, thin rectangular injection ports (thin film fueling). However, the three pylon configurations incorporate different aft shapes to facilitate fuel/air mixing. A cold flow fuel injection study is accomplished to compare mixing capabilities and total pressure losses of the three pylon configurations. The ramp and alternating wedge pylons show decisive increases in mixing capability compared to the basic pylon. They also exhibit a slight increase in total pressure loss compared to the basic pylon.
Cavity flameholders in supersonic combustion ramjet (scramjet) combustors, while effective, fail to take advantage of the full combustor volume. Adding a pylon to the leading edge of a cavity flameholder generates a flowfield increasing mass exchange between the cavity and main combustor flow, increasing the mixing interface between flameholder products and main combustor flow, and exhibiting minimal Reynolds number effects. To demonstrate this modified flowfield driven by supersonic expansion behind the pylon, pylon-cavity flameholder flowfield data were obtained through a combination of wind tunnel experimentation and steady-state computational fluid dynamics. Flowfield effects of the pylon-cavity were examined at a Mach number of two and Reynolds numbers from approximately 33 million m-1 to 55 million m-1. Addition of the pylon resulted in approximately three times the mass exchange between the cavity and overlying flow. Reynolds number effects were weak. A strong upward flow behind the pylon, from the cavity to the top of the pylon wake, significantly increased exposure and exchange of cavity fluid with the main combustor flow. Assuming a suitably reacting fuel-air mixture, the addition of a pylon offers the scramjet designer an attractive option to take advantage of a greater proportion of combustor volume and improve combustor performance.
The size specifications for suitable tracer particles for particle image velocimetry (PIV), particularly with respect to their flow tracking capability, are discussed and quantified for several examples. A review of a wide variety of tracer materials used in recent PIV experiments in liquids and gases indicates that appropriately sized particles have normally been used. With emphasis on gas flows, methods of generating seeding particles and for introducing the particles into the flow are described and their advantages are discussed.
A numerical study of mixing and combustion enhancement has been performed for a Mach 2 model scramjet (supersonic combustion ramjet) combustor. Fuel (hydrogen) is injected at supersonic speed through the rear of a lobed strut located at the channel symmetry axis. The shape of the strut is chosen in a way to produce strong streamwise vorticity and thus to enhance the hydrogen/air mixing. Strength and size of the vortices are defined by the strut geometry and may be modified. It will be shown that in comparison to planar struts the mixing efficiency is strongly improved. On the other hand, the induced vortices cause an increase in entropy and larger losses in total pressure. Different planar and lobed strut injectors are investigated numerically and a comparison with experimental data is given for cold supersonic mixing. Based on this study a numerical investigation of flame stabilization and fuel burnout is performed where two stable modes of combustion are identified. They are associated with attached or detached flames depending on the chosen inflow conditions. In both cases subsonic regions at the channel symmetry axis are responsible for flame holding. If the combustor geometry is chosen in a favorable way these subsonic zones may be kept small. Moreover, the flames are away from solid walls thus minimizing the wall heat load.
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