Supersonic combustion Research Papers

Magnetohydrodynamic Augmentation of Pulse Detonation Engines

2024, Bulletin of the American Physical Society

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Development and characterization of a Fabry-Perot refractometer for gas pressure measurement

2024, HAL (Le Centre pour la Communication Scientifique Directe)

Pour obtenir le grade de : Docteur d'HESAM Université préparée au : Conservatoire national des arts et métiers Discipline : Science des métiers de l'ingénieur Spécialité : Lasers, nanosciences et métrologie Développement et... more

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Large Eddy Simulation of a supersonic lifted flame using the Eulerian stochastic fields method

by Salvador Navarro-Martinez

2024, Proceedings of the Combustion Institute

Scramjet propulsion systems can be the key to deliver the next generation of hypersonic planes. The high costs and complexity of gathering experimental data is a limiting factor in the development of such engine. In this context,... more

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Numerical Simulation of Colorless Distributed Combustion with LES/FMDF

by Ashwani Gupta

2024, Bulletin of the American Physical Society

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Partial equilibrium and relaxation in a plasma in thermal non-equilibrium

by Roger Prud'homme

2024, HAL (Le Centre pour la Communication Scientifique Directe)

Un gaz tel que l'azote peut, lorsqu'il est porté à haute température, mettre en jeu des processus chimiques complexes. Nous avons essayé de caractériser les vitesses de ces processus en calculant les temps chimiques dans différentes... more

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Numerical Analysis of a Kerosene-Fueled Scramjet Combustor

by Pulin Manna

2024, Journal of Thermal Science and Engineering Applications

A kerosene-fueled scramjet combustor was numerically analyzed in order to meet the requirement of thrust for a hypersonic test vehicle. The internal configuration of the fuel injection struts and fuel injection was arrived through... more

Fig. 5 Comparison of nonreacting top wall pressure distribu- tion (Z= 1.6h)

Fig. 3 Grid distribution at different planes of the scramjet combustor: (a) X-Y plane, (b) X-Z plane, and (c) Y-Z plane 4.2 Boundary Conditions. All the flow properties are kept constant at the inflow plane, as the inflow boundary is supersonic. No slip and adiabatic wall boundary conditions are imposed at the walls. The ambient pressure outflow boundary condition has been imposed at the exit of the combustor as the combustor is tested at ambient condition. When the combustor is analyzed to calculate the thrust availability, simulations have been carried out with Due to reaction of the fuel, Mach number values come down, static pressure and temperature increase in the middle of the

Fig. 1 Details of the scramjet combustor: (a) top view and (b) side view

Fig. 4 (a) Mach number, (6) static pressure, and (c) static temperature distribution at various axial planes

Fig.6 (a) Mach number, (b) static pressure, and (c) static temperature distribution for reacting flows at various axial planes Mass fraction of CO3, Oz, and C,2H»3 at various axial locations is shown in Figs. 8(a)—-8(c), respectively. Reaction pattern for both the modules (LM and RM) looks almost similar. Liquid kero- sene droplets completely vaporize within the combustor and no liquid droplets are found at the exit of the combustor. However, some of the unburnt kerosene is observed adjacent to the middle wall regions. Combustion efficiency is defined as the ratio of the

Fig. 7 Average Mach number (a), static pressure (b), and static temperature (c) distribution for reacting flows along the length

Fig.8 Mass fraction distribution of (a) COz2, (b) O2, and (c) Jet-A (C;2H23) vapor at various axial locations

burnt kerosene to the total amount of kerosene fuel injected in the combustor and found to be 78.2% for the present case. Thrust achieved from the combustor is calculated from the difference of the outlet momentum and inlet momentum of the combustor and is found to be 7350 N/kg s of fuel flow rate. Fig. 9 Comparison of top wall pressure distribution (Z= 1.6h)

Fig.10 Effects of vitiated air flow (RM) Acknowledgment

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Raman scattering measurements of mixing and finite-rate chemistry in a supersonic reacting flow over a piloted, ramped cavity

by Robert Pitz

2024, Combustion and Flame

UV Raman scattering is applied to measure fuel/air mixing and combustion of a Mach-2 air stream flowing over a step-ramp cavity fueled with 70% methane/30% hydrogen. Average and RMS fluctuations of temperature and major species profiles... more

Figure 1: Research Cell 19 wind tunnel schematic (top), and close-up of combustor and cavity (bottom) with the blue dot located on the leading edge of the cavity and the centerline of the combustor (x = 0, y= 0, z=0)

Figure 4: Sample mean wall pressures and the standard deviation of the 1-Hz measurement was generally <1%. Insert shows the pressure port locations (solid triangles) with respect to the cavity. Flow pattern sketched following Figure 5b in [20].

Figure 7: Calibration factors, C(T), vs. temperature obtained from the Hencken burner. The C(T) for CH, measured at room temperature and assumed to be invariant with temperature.

Figure 8: Raw, single-shot line spectral images of calibration data of 70% CH4/30% Hh) fuel mixture at @ = 1 (left) and o = 1.6 (right).

Figure 21: Scatter plots for Location 1 (x = 3.5 cm, y = 1 cm) with lines drawn for adiabatic equilibrium (solid) and mixing without reaction (dashed).

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Fluid Dynamic Analysis Over Turbulent Flows

by Tevin James

2024

The purpose of this research is to run CFDs (Computational Fluid Dynamics) with several type of objects that consist of the following: 90-Degree Pipe, Combustors, Scramjets, Scramjets with Fuel Injectors, and Airfoils. The sole reasoning... more

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Dual Injection in Supersonic Crossflow: Analysis Jet Shear Layer from Schlieren Images

by Harry W . M . Hoeijmakers

2024, AIAA Journal

In supersonic-combustion ramjets (scramjets), fuel is injected, which should mix rapidly with the supersonic crossflow to minimize the length of the scramjet. Tandem dual-jet injection has shown improved mixing performance over single-jet... more

Fig. 2 Schematic of tandem dual sonic jet injection into supersonic crossflow. Adapted with permission from Ref. [3] (copyright 2006 by ATAA). Fig. 1 Schematic of a scramjet in which flow is supersonic throughout whole engine. Adapted with permission from Ref. [2] (copyright 2020 by AIAA).

Fig.3 Supersonic wind-tunnel facility with schlieren setup at University of Twente. Adapted with permission from Ref. [2] (copyright 2020 by ATAA). The inlet of the wind tunnel (left) is at atmospheric conditions, providing the total temperature T),. and the total pressure po... Perpendicular to the bottom wall, pressurized air is injected at sonic speed from two (tandem) orifices. The air in the jet(s) originates from a plenum in which the total pressure of the jet flow is measured using a single plenum GE Druck DP 104 pressure sensor. For tandem dual-jet injection, air is injected into the crossflow through two orifices with diameters D; = 1 mm and D, = 2 mm, respectively. This results in D,/D, = 0.5 and a total diameter of D, = /D{ + D3 = J5 mm. For single-jet injection, air is injected through a single orifice, with diameter D = D, = V/5 mm, chosen to be equal to the total diameter

Fig. 4 Schematic of side view test section supersonic wind tunnel at University of Twente [5]. The test section, marked by dashed lines, is slightly divergent, which corrects for effect of vertical boundary layer.

Fig.5 Schematic overview of present schlieren setup. The left and right walls of the test section are glass windows.

Fig.6 Features of tandem dual sonic jet injection into supersonic cross- flow (J = 3.8, S = 4.48, and M, = 1.55), shown in an unprocessed schlieren image.

Fig.8 Flow features in jet plume as determined from a schlieren image. Features in jet are indicated by clusters of closely spaced blue asterisks (J = 3.8, S = 4.93, and M, = 1.55). For each schlieren image, the flow features within the jet plume are captured as sets of coordinates. From the sets of coordinates of the clusters in a single image (see, e.g., Fig. 8), for every x coordinate, the corresponding y coordinates of the upper and lower edges of the plume are determined and stored as two additional sets of coordi- nates. Figure 9 illustrates the result for the upper and lower edges of the jet plume. Each image provides +350 data points for the upper edge and for the lower edge of the jet plume. Usually, the jet upper shear layer can be determined satisfactorily; however, because of the mushroom shape of the cross-section of the jet in the plane normal to the tunnel axis and because of the vortices forming in the jet plume, the location of the lower edge of the jet plume is not always well defined. For more details about the lower edge of the jet plume, see the work by Smink [17]. For each (J, S) combination, the sets of coordinates of all 20 schlieren images recorded are combined and stored for further analysis.

Fig. 7 Window (binary mask) determined within schlieren image. The window covers up areas not relevant for analysis of behavior of jet, areas that otherwise might disturb quality of results of postprocessing. (J = 3.8, S = 4.93, and M, = 1.55).

Fig.9 Upper (blue) and lower (red) edge of jet plume upon application of postprocessing for a single schlieren image (J = 3.8, S = 4.93, and M, = 1,55).

Fig.10 Isoincidence plot obtained from 20 schlieren images. Features inside jet are visualized as dark-red islands of higher incidence, which represen stationary, or frequently occurring, features (J = 3.8, S = 4.93, and M, = 1.55). The schlieren images show that for specific combinations of J and S the two bow shocks merge. This point is called the merger point. The distance between the two bow shocks as well as whether or not the two bow shocks merge directly depends on J and S. The position of the merger point also depends on J and S; however, it is noted that the location of the merger point fluctuates due to significant time variation of the location of the downstream bow shock. The oscillatory character of the bow shocks was already observed in Refs. [9,18]. Gruber et al. ‘Figure 6 shows that in front of each of the two jets from the bottom wa the of t 1a bow shock is formed, which is reflected at the opposite wall of test section. The position of the jet upper shear layer downstream he second orifice is considered indicative for the penetration of the jet into the crossflow. It is also a measure for the extent the gas from the primary jet mixes with the gas from the crossflow. Note that the slope of the Mach lines, which are caused by small irregularities at the wind-tunnel walls, are a measure for the Mach number of the undis- turbed crossflow: sin(w) = 1/M_,, with y the slope of the Mach line wit h respect to the horizon.

Fig. 11 Unprocessed schlieren images showing effect of variation in S at constant J = 3.8. Bow shocks merge closer to jet orifices for smaller S. Also, bow shocks merge closer to jet orifices for larger J. When bow shocks are sufficiently separated, then the two bow shocks behave independently. However, when they approach each other, then downstream bow shock changes shape. a) J = 3.8, S = 3.59, and M, = 1.55. b) J = 3.8, S = 4.48, and M, = 1.55. ce) J = 3.8, S = 5.38, and M, = 1.55. d) J = 3.8, S = 9.87, and M, = 1.55.

Fig. 12 Determination of location merger point from 20 images. The merger point is indicated by blue asterisks. The red asterisk indicates the time-averaged position of merger points (J = 3.8, S = 4.93, and M.. = 1.55). Remarkably, the location of the merger point is always found on the upstream bow shock, with the precise position depending on the specific (J, S) combination. The location of the merger point is found to vary within a certain range around its time-averaged location (€,n) = (x/D; + S, y/D,) which is indicated by the red marker in Figs. 12 and 13 (range of [—1.90D,, 1.83D,] along the upstream bow shock, with variance of 0.567D,). This excursion around the time- averaged location is due to the fluctuation of the downstream bow shock (see Fig. 12). The origin of the coordinates (&, 7) is located at the center of the upstream orifice so that € = S is at the center of the downstream orifice.

Fig.13 Position of merger point for all (J, S) combinations (x), and their mean position per (J, S) combination (@). (M, = 1.55). Included are predictions of Eq. (4) for J € [2.8,3.8,4.8].

Fig. 14 Instantaneous location of upper shear layer jet from schlieren images (blue) and empirical similarity relation for time-averaged location (Eq. (8), black), both scaled with c,(J,S) [see Eqs. (6) and (8)] for J = [2.8,3.8], S € [0:9.87], M, = 1.55. Data obtained from 20 schlieren images per (J, S) combination, total of 22 x 20 images. Each schlieren image provides +350 data points. The coefficient of determination of R? = 0.739 calculated for this fitting function, calculated for all data acquired, is considered

Fig. 15 Instantaneous location of upper shear layer jet from schlieren images (blue) and empirical similarity relation for time-averaged location [Eq. (8), black] for J = 4.8, S = 5.38, and M, = 1.55. Data obtained from 20 schlieren images per (J, S) combination.

Fig. 16 Bar diagram: spatially and time-averaged penetration depth of upper shear layer of jet as calculated from individual least-squares fits using Eq. (6). Solid lines: spatially and time-averaged penetration depth calculated from Eq. (9). Optimal value (x) of S at constant J for maximal penetration depth equals S,,; = 4.732J°7*8 where yay./D, equals 3.582J°*!, M, = 1.55. Other questions that remain are as follows. What is the difference between penetration for S = 0 and that for § > co? What is the behavior of the upper shear layer of the jet for S > 10, and which to be acceptable; see also Fig. 14. The variance calculated from the data (x/D,;, y/D,) and Eq. (8) is equal to 0.352. Using Eq. (7), the average penetration depth calculated from the empirical similarity relation is

Fig.17 Bar diagram: spatially and temporally averaged penetration depth in domain (0.5 < x/D, < 15) from experiments by de Maag [2]. Solid lines: average penetration depth calculated from Eq. (9). The maximal penetration depth (x) is calculated from S,,; = 4.732J°88, where Yavg/D; equals 3.582941, M, = 1.6.

Fig.A1 Unprocessed schlieren images for J = 2.8 and all values of S considered. M, = 1.55. For every (J, S) combination, one of the 20 images available is shown.

Fig. A2 Unprocessed schlieren images for J = 3.8 and all values of S considered. M, = 1.55. For every (J, S) combination, one of the 20 images available is shown.

Fig. A3_ Unprocessed schlieren images for J = 4.8 and the two values of S considered. M, = 1.55. For every (J, S) combination, one of the 20 images available is shown.

Table 1 Conditions in test section for injection into supersonic crossflow (left) and in jet (right) (T, po, and M are measured, and all other quantities are calculated from isentropic relations for calorically perfect gases)

Table 2 Results for c,; determined from least- squares fit, compared with results from Eq. (8), together with coefficient of determination R?, for J = 48, S € [0,5.38], and M, = 1.55

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Effect of the Location of an Incident Shock Wave on Combustion and Flow Field of Wall Fuel-Injection

by Hideaki Kobayashi

2023, TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES

Experimental and numerical studies on the interaction between combustion of a hydrogen jet and an incident shock wave were performed. When the incident shock wave was introduced upstream of the injection slot, the boundary-layer... more

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In-flight imaging of transverse gas jets injected into subsonic and supersonic crossflows

by Ann Karagozian

2023, 33rd Aerospace Sciences Meeting and Exhibit

The development of and results from an airborne flight-test experiment to study nonreacting gas jets injected transversely into subsonic and slightly supersonic cmssflows is presented. Freestream/crossflow Mach numbcrs range fmm 0.6 to... more

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Experimental and Theoretical Investigations of a Triple Vortex Combustor with Spatial Arc

by Serhiy Serbin

2023, 46th AIAA Aerospace Sciences Meeting and Exhibit

Theoretical and experimental investigations of working processes in a triple vortex combustor with spatial arc have been conducted. Selected concept of portable triple vortex plasma assisted combustor can provide higher performance, wider... more

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Quadrant Analysis of a Mixer-Ejector Nozzle for Supersonic Transport Applications

by Nicholas Georgiadis

2023, Journal of Propulsion and Power

NOTES are short manuscripts describing new developments or important results of a preliminary nature. These Notes should not exceed 2500 words (where a figure or table counts as 200 words). Following informal review by the Editors, they... more

Fig. 1 Schematic of downstream mixer (DSM) nozzle.

Fig. 2. Comparison of internal flow computational domains for quadrant and half-chute calculations.

Fig. 3 Axial velocity contours of quadrant solution inside mixing duct. nozzle calculations. However, the grid topology and density for the half-chute and quadrant simulations were selected based upon previous analysis performed by the multiorganizational HSR CFD analysis team [6].

Fig. 5 Cross-flow stream lines within the mixer—ejector nozzle. Further examination of the quadrant solution was undertaken to better understand the source of the low-energy flow regions along the sidewalls. It was originally speculated that, as the freestream spilled over the external sidewall and into the secondary inlet, it rolled into a vortex. This vortex was then thought to pass through the bottom of the secondary chute nearest the sidewall and along the length of the mixing duct. Examination of the streamlines in Fig. 5 indicates that a Convergence criteria for these cases included monitoring of the residual error as well as integral flow quantities such as mass flow rates through the mixing duct, thrust coefficients, and pumping ratio. Solutions were determined to have reached steady state once a three- order reduction in residual had been achieved and the mass flow rate through the mixing duct was conserved within 0.5%. The resultant

Table 1 Distribution of grid points within the computational domain

Table 2 Nozzle performance results values for the pumping ratio and gross thrust coefficient were converged within 0.2% and 0.3%, respectively.

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Kinetic Modeling of Cyclohexane and n-Propylcyclohexane Oxidation with the PAH Precursor Formation

by Uwe Riedel

2023

A reaction mechanism of cyclohexane (cyC6H12) and n-propylcyclohexane (cyC9H18) is developed to study its oxidation at both low and high temperatures, including PAH precursors routes. The cyclohexane oxidation kinetic mechanism is a... more

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A Smart Simple Virtual Instrument for Measuring the Electrical Signal

by ronaldo talapessy

2023, SCIENCE NATURE

Real-time measurement, especially in analyzing the electrical waveform, is necessary to explore physical phenomena. Virtual instrumentations have been studied for various applications such as water existence, characterization and modeling... more

issociated with the determination of resistivity of materials in the LabVIEW provides many kinds of libraries that enable us to develop GUI programs that operate to measure electronic equipment and analyze the signal. LabVIEW programs have been developed for designing virtual digital oscilloscopes, real-time data acquisition for tomography, plasma _ diagnostic, scanning micro-beam X-ray fluorescence spectrometer, monitoring, and simulation system for photovoltaic systems [1-5]. In addition, the LabVIEW program has been applied to monitor and calculate the electrical properties of the ground in the field measurement [6, 14]. abVIEW environment to monitor and analyze the square wa sraphical User Interface (GUI) has been programmed in a

III. Software demonstration A Smart Simple Virtual Instrument for Measuring the Electrical Signal

Figure 3. Block diagram of serial USB: (a). Connect USB, (b) Disconnect USB. Figure 2. The front panel of the set interface.

oscillations made within a specific time. dM Order to test the system, we Current Was injected into the soil sample through a couple of electrodes that show on the front panel. Figure 4a, 4b, and 4c present the measurement of soil samples with different frequencies, such as 33 Hz, 38 Hz, and 40 Hz, respectively. Voltage and current were monitored on the front panel as a function of time (ms), where frequency measures the total number of oscillations made within a specific time. Figure 4. The front panelof the virtual instrument for various frequencies: (a) 33 Hz, (b) 38 Hz, (c) 40 Hz. Figure 4. The front panelof the virtual instrument for frequency measures the total number of various frequencies: (a) 33 Hz, (b) 38 Hz, (c) 40 Hz.

Figure 5. Block diagram of virtual instruments. the total amount cut equals twice the value of

Figure 6. The front panel of earth resistivity simulation.

Figure 7. Block diagram of simulation. iterative measurement function, a for loop shown on the chart of the front panel. As an

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The effect of injection angle on mixing and flame holding in supersonic combustor

by Mohamad Ali

2023, Journal of Thermal Science

A numerical study on mixing of hydrogen injected into a supersonic air stream has been performed by solving two-dimensional full Navier-Stokes equations. An explicit Harten-Yee Non-MUSCL Modified-flux-type TVD scheme has been used to... more

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Numerical Simulation of Mixing and Combustion of a Hydrogen Fueled Scramjet combustor using Strut Injectors

by SIVA SAKTHI

2023

A major problem in supersonic combustion is the short residence time of the fluid inside the combustor due to high flow velocities. Thus techniques for mixing enhancement have to be used to achieve a fast and efficient fuel-air mixing. In... more

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Combustion and Mixing Analysis of a Scramjet Combustor Using CFD

by Dr P Ravinder Reddy

2023, International Journal for Scientific Research and Development

The present study is to determine flow field in the three-dimensional scramjet engine combustor with coupled implicit NS equations, the standard k-e turbulence model are used and the finite-rate/eddy-dissipation reaction model has to be... more

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Fluid Dynamic Analysis Over Turbulent Flows

by Dareon Medley

2023

The purpose of this research is to run CFDs (Computational Fluid Dynamics) with several type of objects that consist of the following: 90-Degree Pipe, Combustors, Scramjets, Scramjets with Fuel Injectors, and Airfoils. The sole reasoning... more

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Simulation of a Scramjet Combustor: A Priori Study of Thermochemistry Tabulation Techniques

by Lisa Bouheraoua

2023, Flow, Turbulence and Combustion

Large eddy simulations (LES) of a scramjet combustor are reported in this paper. The case under study is a cavity-based combustion chamber that is experimentally studied at the U.S Air Force Research Laboratory. The chamber is fed by... more

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Numerical characterization of 3D nonreacting supersonic cavity combustor with inlet Mach number variation

by Yared Mengistu

2023, International Journal of Hydrogen Energy

h i g h l i g h t s A 3D simulation has been carried out to study for a cavity-based supersonic combustor. A very complex flow structure with horseshoe vortices and counter rotating vortices is observed. The performance characteristics... more

The flow field to be analyzed utilizes a three-dimensional, compressible, turbulent time-averaged Navier Stokes equa- tions by neglecting body forces and heat source terms. The

Fig. 2 — Computational domain of cavity-based supersonic combustor.

Fig. 3 — Grid Independence study for model combustor.

Fig. 4 — Static pressure distribution along the lower wall center-line.

Fig. 5 — Pressure distribution at the leading and trailer edge of the cavity for Ma = 2.5.

Fig. 6 — Pressure contour for displaying different forms of shocks for a) Ma = 1.5, b) Ma = 2.5, c) Ma = 3.5. (BS: Bow Shock, RS: Reattached Shock, EF: Expansion Fan, SR: Reflected Shock).

Fig. 7 — Density contours on X—Y plane a) Ma = 1.5 b) Ma = 2.5 c) Ma = 3.5.

Fig. 8 — Velocity vector plot on x-y plane a) Ma = 1.5 b) Ma = 2.5 c) Ma = 3.5.

Fig. 9 — Pressure contour for Ma = 2.5 at different planes exhibiting 3D forms of shock waves (Y = 0, X = 0, 30 mm and Z = 25 mm).

Fig. 10 — Hydrogen mass fraction for Ma = 1.5 on Y-Z plane at x = 0.0 and x = 200 mm. recirculation zone. In the cavity, the extent of mixing gets enhanced relatively for all cases of Mach number due to the increased residence time of fuel. Besides this, the fuel injector is relatively closer to the cavity fore-wall in the present combustor model. As a result, the injected fuel turns due to the presence of cavity configuration before penetrating the incoming the freestream flow; so that significant part of the fuel-injected enters into a cavity. Subsequently, the fuel in this cavity gets thoroughly mixed with air due to the forma- tion of a recirculation zone where the fuel and air recirculate The free-stream Mach number also impacts the fuel entrainment and its distribution in the vicinity of cavity. Figure 13 exhibits distribution hydrogen mass fraction inside the cavity in the x-z plane (y = -2, and -10). It can be noted that the fuel jet gets deflected to strike the angled rear wall of the

Fig. 11 — Hydrogen mass fraction for Ma = 3.5 on Y-Z plane at X = 0.0 and 200 mm.

Fig. 12 — Average Mass fraction of H2 on Z-Z plane for a) Ma = 1.5 and b) Ma = 3.5 at X = 30, 60 mm, 90 mm, 120 mm, 150 mm, 200 mm. cavity by the reattachment wave and subsequently, flows back to fill the cavity caused due to the formation of the recirculation zone. However, during this backflow phenome- non, uneven distribution of the fuel is observed in the cavity especially for higher Mach number cases (Ma = 2.5 and 3.5). Interestingly, Ma = 1.5 exhibits a uniform fuel distribution within the cavity region (Fig. 13a). The z-x plane on the bottom wall of the cavity at y = - 0.8 mm shows symmetric Further, a similar study is conducted on x-y plane to investigate the fuel distribution in the cavity and its down- stream region (Fig. 14). The fuel distribution in the center plane (z = 25 mm) reveals very little entrainment of the

Fig. 13 — Fuel distribution on lower and upper plane cavity Z-X plane at Y = —0.2 and —10 mm for a) Ma = 1.5 and b) Ma = 3.5.

mainstream fuel into the cavity region for all Mach number cases. However, the region away from the centerline plane (z = 35 mm) shows an enhanced extent of fuel entrainment into the cavity as shown in Fig. 13b. A similar effect, but with different magnitude of fuel mole-fraction, is observed for all Mach number cases. Hence, the three-dimensional effect plays a vital role in the characterization of air-fuel mixing in supersonic cavity-based combustors, in addition to the rea- sons discussed above. Fig. 14 — Fuel distribution on Y-X plane at a) Z = 25 and b) Z = 35 respectively for Ma = 3.5. In order to compare the behavior of the combustor in cold flow conditions for various inlet Mach number the performance parameters namely, the coefficient of pressure, total pressure loss and mixing efficiency are considered for the present study.

Fig. 15 — Coefficient of pressure at lower wall center line. Please cite this article as: Mengistu YG et al., Numerical characterization of 3D nonreacting supersonic cavity combustor with i: Mach number variation, International Journal of Hydrogen Energy, https://doi.org/10.1016/j.ijhydene.2020.01.187 The variation of total pressure loss downstream of the fuel injection starting from the cavity leading edge for all three The mixing efficiency for the present work is calculated on y-z pane at a certain constant x-distance downstream of the fuel

Fig. 16 — Mixing efficiencies for three Mach number cases.

Fig. 17 — Total pressure loss for three Mach number cases. Mach number cases are plotted against the axial distance, as shown in Fig. 17. It can be observed that total pressure loss for higher Mach number (Ma = 2.5 and 3.5) is larger than that of the lower Mach number case (Ma = 1.5). The total pressure loss is observed to be highest in the case of Ma = 3.5 and lowest in the case of Ma = 1.5 for a constant injection pressure upstream of the cavity. The pressure loss for Ma = 2.5 and Ma = 3.5 cases attain almost a similar value inside the cavity due to non- uniform recirculation pattern, but at the downstream of the cavity, the pressure loss for the higher Mach number is The effects of inlet Mach number on the flow structure and air- flow mixing phenomena on a cavity-based combustor with upstream fuel injection under non-reacting flow conditions are investigated numerically by considering three different

Table 1 — Supersonic combustor dimensions. turbulence is modeled using a two-equation SST x-w model with the estimation of density through the ideal gas law.

Table 2 — Angles of reattachment and reflected shocks for all Mach numbers. Please cite this article as: Mengistu YG et al., Numerical characterization of 3D nonreacting supersonic cavity combustor with i1 Mach number variation, International Journal of Hydrogen Energy, https://doi.org/10.1016/j.ijnydene.2020.01.187

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Studies of a combined way of flame stability in ramjet combustor

by Jiaxin Tao

2023, Defence Technology

A numerical simulation was conducted to study the influence of bleeding. The Euler-Lagrange method was used to investigate the two-phase turbulent combustion flow. Standard k-ε turbulent model was adopted in the continuous phase... more

so that even if the flow condition is poor, the stable piloted flame can also be produced. rapped vortex can be produced in the inner region of cavity flame holder, which makes the energy in the

High temperature air enters from the bottom of the strut and then mixes with oil which is ejected from the The combustor model used in this paper is shown in Fig. 2(a). The combustor is rectangular cross section, and the Computational Model

produced after air bleeding. According to the trapped vortex combustion theory, the trapped vortex combustion can occurs in the region directly behind the bluff body, and the combustion of the cross flow direction is not ideal. Until the exit of the nozzle, there is still a large area with almost no CO, distribution. A fter air bleeding, combustion in a short distance to realize the diffusion of cross flow, inlet CO, distribution is uniform, the air bleeding plays a certain role in transverse cross flame. through the trapped vortex. Ultimately the flame spreads along the axial direction with the mainstream. In the

Because the influence of air bleeding condition on combustion performance is not clear, it is necessary to

Fig. 6. Curves of total pressure recovery coefficient to bleeding temperature Because the reaction used here is a single step overall reaction, the calculation formula of combustion efficiency

It can be seen from the Fig. 6 that the air bleeding has almost no influence on the total pressure before the strut. When the air bleeding is applied, due to the shear effects on the mainstream, the total pressure recovery coefficient along the axial direction is slightly lower than the no air bleeding state. In the nozzle exit total pressure recovery coefficient is reduced by about 0.02. Change of bleeding temperature has little effect on the total pressure recovery It can be seen from the Fig. 6 that the air bleeding has almost no influence on the total pressure before the stru coefficient is reduced by about 0.02. Change of bleeding temperature has little effect on the total pressure recovery

Fig. 9. Curves of combustion efficiency to gross fuel-air ratio Fig. 8. Curves of total pressure recovery coefficient to gross fuel-air ratio

Fig. 8. and Fig. 9. are the curves of total pressure recovery coefficient and combustion efficiency along the axia direction under different fuel-air ratio. Total pressure recovery coefficient near the flame holder is lower and the total pressure recovery coefficient is higher than anywhere else, which indicates that through air bleeding the flame holder can make greater resistance and lower mixing loss in a lean oil state. The combustor combustion efficiency a the outlet reaches 0.98 when the gross fuel-air ratio is 0.03, which is much higher than when the gross fuel-air ratic Fig. 8. and Fig. 9. are the curves of total pressure recovery coefficient and combustion efficiency along the axia is more difficult, this is the main reason leading to the combustion efficiency decrease. holder can make greater resistance and lower mixing loss in a lean oil state. The combustor combustion efficiency at

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Experimental Study on Effects of Shock Wave Impingement on Supersonic Combustion

by Nur Hakim

2023, Memoirs of the …

An experimental study has been conducted using reflected type of shock tunnel to investigate the effect of shock wave impingement on supersonic combustion. In the experiment, air is compressed by reflected shock wave up to total... more

Fig.1 Schematic Diagram of Shock Tunnel Facility.

Fig.2 Test Section (nozzle wall is inclined by 1 degree from horizontal axis x).

The total pressure is obtained in the low pressure section by measuring pressure of test gas at the stagnation condition after the incident shock wave is reflected. Measurement of this pressure is made by using Kistler pressure transducer located 10 mm before nozzle throat. On the other hand, the measurement of the stagnation heat flux of semi-sphere cylinder model, which is located at the center of test section and at the same position with injection point, has been conducted without injection of gas in order to obtain the total temperature of test air. The measurement is conducted by measuring temporal surface temperature rise at the stagnation point of semi-sphere model. For the measurement of surface temperature thin film platinum heat transfer gauge is used. Assuming an isentropic expansion of test air from the driven section to the test section and using combined relationship between total pressure of reservoir and the heat flux in the test section, the total temperature of test air in the test section can be obtained. the test section, the total temperature of test air in the test section can be obtained.

Fig.5 Time history of total temperature of freestream calculated from heat flux and total pressure. ig.5 Time history of total temperature of freestream in the test section.

Fig. 8 Enlarged image of shock structure around interference point (Ramp angle 15°); Helium injection (left) and Hydrogen injection (right).

In the case of ramp angle of 10° ( Fig. 9 (b) ) the reaction starts just behind the oblique shock wave. The reaction becomes stronger after the impingemen of the second oblique shock wave, which is formed at the bottom wall just after the ramp. In the case of ramp angle of 10° ( Fig. 9 (c) ) very strong combustion occurs just behind the oblique s hock wave. In this case, the region of combustion is wider than the other two cases. The incident shock wave interacts with the boundary layer on the upper wall and generates Mach stem and reflected shock wave. Also at the foot of Mach stem the boundary layer separates and the recircu ation region is formed. As shown in the Schlieren photograph and UV image, just after the front s and combustion zone becomes wider. hock wave hydrogen begins to burn

Table 2 Change of Flow Properties across the Oblique Shock (M\=2).

Table 3 Experimental Conditions. 2.3 Flow visualization

Table 4 Comparison of the theoretical and experimental shock wave angle. Planar oblique shock waves are generated by ramp wall mounted at the bottom wall of the test section and interact with the air-fuel mixture near upper wall and then reflected by upper wall. The angles of shock wave generated by each ramp both in reacting case and non-reacting cases are listed in Table 4. As seen in the table, the angle of shock wave generated by the ramp slightly deviates from its theoretical value. This change of shock angle causes the shift of the impingement point of the oblique shock.

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Numerical simulation and validation of high-speed turbulent and chemically-reacting flows

by Mohammed Kamel

2023, International Journal of Engineering Systems Modelling and Simulation

Mixing and combustion processes in scramjet engines involve complicated aerothermochemical features such as interactions between shock-waves and boundary-layer, shock induced-combustion and recirculation zones. In this study, a numerical... more

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Numerical simulation of supersonic and turbulent combustion with transverse sonic fuel injection

by Mohammed Kamel

2023, International Journal of Engineering Systems Modelling and Simulation

Complicated interacting flow features such as shock-wave, boundary-layer and shock induced combustion are simulated numerically in the current study to investigate the effect of the transverse sonic fuel injection on the air-fuel mixing... more

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Visualization of Oblique Shock Waves Incident on Pointed Surfaces with Ramps

by Edna Santiago Benta

2023, Journal of Engineering Research

All content in this magazine is licensed under a Creative Commons Attribution License. Attribution-Non-Commercial-Non-Derivatives 4.0 International (CC BY-NC-ND 4.0).

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The application of supplementary fuels for the control of supersonic reacting air-fuel mix flows in the combustion chamber

by Stanislav Kolosenok

2023, Technical Physics Letters

Besides gas-dynamic methods, chemical ones are also suitable for the implementation of stable supersonic combustion of hydrocarbon fuels. Organoelemental compounds are known for their high reactivity, so attention was paid to... more

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Mixing and Combustion Analysis of a Hydrogen Fueled Scramjet Engine Using Strut with Alternating Wedge Injector

by Surya Kumar

2023

A major problem in supersonic combustion is the short residence time of the fluid inside the combustor due to high flow velocities. Thus techniques for mixing enhancement have to be used to achieve a fast and efficient fuel-air mixing. In... more

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Détermination de constantes de vitesse de réactions chimiques élémentaires importantes dans la combustion de l'hydrogène en régime d'écoulement supersonique

by Sandra JAVOY

2023, Http Www Theses Fr

Je tiens à exprimer ma profonde gratitude à M. Claude-Etienne PAILLARD, professeur à l'Université d'Orléans, pour avoir suivi et dirigé cette thèse, m'avoir fait confiance et pour ses encouragements. J'ai pu profiter de l'étendue de ses... more

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Experimental and Theoretical Investigations of a Triple Vortex Combustor with Spatial Arc

by Serhiy Serbin

2023

Theoretical and experimental investigations of working processes in a triple vortex combustor with spatial arc have been conducted. Selected concept of portable triple vortex plasma assisted combustor can provide higher performance, wider... more

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Numerical Analysis of Integrated Liquid Ramjet Engine

by Raja Singh Thangadurai

2023, Defence Science Journal

The numerical simulation of an integrated, liquid-fuelled ramjet engine comprising supersonicair intake, subsonic combustor and a convergent-divergent nozzle has been carried out and theresults are discussed in this paper. These results... more

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Tip-to-tail numerical simulation of a hypersonic air-breathing engine with ethylene fuel

by debasis Chakraborty

2022, Acta Astronautica

End to end CFD simulations of external and internal flow paths of an ethylene fueled hypersonic airbreathing vehicle with including forebody, horizontal fins, vertical fins, intake, combustor, single expansion ramp nozzle are carried out.... more

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Experimental Study of Propane-Fueled Pulsed Detonation Rocket

by Frank Lu

2022, 12th AIAA International Space Planes and Hypersonic Systems and Technologies

A major problem applying detonations into aero-propulsive devices is the transition of deflagration and weak detonation into CJ detonation. The longer this transition, the longer the physical length of the engine must be to facilitate the... more

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Investigation of flow characteristics in supersonic combustion ramjet combustor toward improvement of combustion efficiency

by Prasanth P Nair

2022, International Journal of Energy Research

Supersonic combustion ramjet (scramjet) is a variant of ramjet in which the combustion takes place at supersonic velocity. The flow physics inside the scramjet combustor is quite complex due to the fact that the mixing and completion of... more

Supersonic combustion ramjet (scramjet) is a variant of ramjet in which the combustion takes place at supersonic velocity. The flow physics inside the scramjet combustor is quite complex due to the fact that the mixing and completion of the combustion take place in a short time, which is of the order of milliseconds. This study focuses on flow characteristics within the combustion chamber of the scramjet engine that is designed to improve energy efficiency by enhancing combustion efficiency. The effect on combustion performance and thereby the energy efficiency on using strut-based flame stabilizer is evaluated at different positions. Reynolds averaged Navier-Stokes equations are solved with the Shear Stress Transport k-ω turbulence model. Single strut configuration is used to validate with the experimental data. Single strut is then compared with three-strut configuration. In the three-strut configuration, the location of the primary strut is kept constant, and the secondary struts are relocated in x and y directions. Combustion performance was evaluated for the cases of flow from primary strut only and through three struts. It was found that the placement of secondary strut in a three-strut configuration plays a vital role in improving energy efficiency. A maximum of 33.86% improvement in combustion efficiency was observed in comparison to the single strut combustor. A reduction in unburned fuel was observed, making the system more energy efficient. If the struts are not placed optimally, the combustion performance of the combustor was observed to be lower than that of a single-strut configuration. The shock reflection and expansion fans within the primary combustion zone and the secondary strut region enhance the combustion efficiency. The wall static pressure was observed to increase with the addition of secondary struts. For certain strut configurations, flow separation was seen on the combustor walls. If the secondary strut was placed close to the primary strut, combustion efficiency was found to enhance. It was seen that

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Heat transfer modes in supersonic hydrogen combustion

by Cesar Olmedo

2022

A miniature wind tunnel has been built which harnesses the power to hold supersonic flows and supersonic combustion. Experiments have been performed to test the sustainability of hydrogen combustion in supersonic Mach flows. Supersonic... more

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Effect of geometric parameters on the drag of the cavity flameholder based on the variance analysis method

by Derek Ingham

2022, Aerospace Science and Technology

Wall cavities have been widely used as flameholders in scramjet engines to prolong the residence time of the fuel and the air in supersonic flow. These devices improve the combustion efficiency of the scramjet combustor, and also impose... more

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Numerical Simulation of Mixing and Combustion of a Hydrogen Fueled Scramjet combustor using Strut Injectors

by siva sakthi

2022

A major problem in supersonic combustion is the short residence time of the fluid inside the combustor due to high flow velocities. Thus techniques for mixing enhancement have to be used to achieve a fast and efficient fuel-air mixing. In... more

the flow is slowed to subsonic levels before it enters the combustor. Fig.1 shows a basic generic Scramjet design. It works by injecting fuel (typically hydrogen) into a flow of supersonic air. The air is at sufficiently high temperature and pressure for the fuel to combust, and the resulting mixture is expelled from the engine at a higher pressure. The Scramjet is composed of four main sections: the inlet, isolator, combustor and exhaust nozzle. These sections can be seen in Fig.1

Fig.2 Shock and Expansion Waves The most important consideration with supersonic flow is that the flow is compressible. A compressible flow is one for which the density cannot be considered constant (for flow below M = (0.3 the fluid can be considered to have a constant density). Compressibility leads to two phenomena unique to supersonic flow - shockwaves and expansion waves Fig.2. There are two types of shockwaves: oblique and normal. Oblique shockwaves and expansion waves are generated when a supersonic flow changes direction — a shockwave when the flow converges and an expansion wave when the flow expands.

Fig.5 Schematic overview of the supersonic combustion chamber [44] all relevant physics of the scramjet combustor. The geometry considered in this project work is the same as the one considered by Oevermann [44]. The computational configuration is simplified in the sense of, all the gambit modelings are done in 3D and fluent analyses are done in 2D. The reason for this simplification is that to reduce the computational time as much as possible, but still retaining all relevant physics of the scramjet combustor.

Fig.6 Scramjet combustion chamber and strut with circular injector A strut with circular injector is placed in the combustion chamber downstream of the nozzle. The height of the 32 mm long strut is 6 mm and diameter of the injector is 2.23mm.

11.2.2 Computational Configuration of Strut with Planer Injector

Fig.11 Strut with alternating wedge injector Fig.10 Scramjet combustion chamber and strut with alternating wedge injector

The grid independence test is carried out for analyzing the effect of grid number on the bottom wall static pressure with combustion chamber length. As illustrated in Fig.4.11, a good agreement is observed between 4, 00,000 cells and 3, 40,840 cells. Hence, the present analysis is carried out using a minimum grid size of 3, 40,840 cells. 11.6 NUMERICAL SOLUTION

Fig.12 Mesh generation in 2D combustion chamber

Comparing the mixing efficiency of the planer injector with circular injector we can observe that there are some changes in the flow field. In circular injector the mixing are over at x=225mm but in planer injector the mixing is over at x =160mm. from this it is clear that strut with planer injector giving good mixing efficiency. observe that there are some changes in the flow field. In circular injector the mixing are over at

Table 1 Inflow conditions of the air stream and hydrogen jet. 11.4 Mesh Generation

It appears from Fig.5.34 that combustion is complete at x ~ 0-170m. Contours of the mass fraction of H20 are given in Fig.5.32 and Fig.5.33. It can be seen that the maximum value for the mass fraction of H2O is reached at a distance of 0-165m, which is just downstream of the wedge However, the distribution of the mass fraction of H2O continues to change till 0-290m, beyonc which it remains the same.

14. STRUT WITH ALTERNATING WEDGE INJECTOR

Fig.20 Pressure Variation at the Bottom and Top Walls

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Spiking of Hydrocarbon Fuels with Silanes-based Combustion Enhancers

by Hartwig Ellerbrock

2022, TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN

The concept of spiking hydrocarbon fuels such as kerosenes with liquid silicon hydrides in order to render the fuel combination hypergolic and to improve the combustion efficiency is presented and preliminarily analyzed. In view of... more

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Thermochemical exploration of hydrogen combustion in generic scramjet combustor

by debasis Chakraborty

2022, Aerospace Science and Technology

This article appeared in a journal published by Elsevier. The attached copy is furnished to the author for internal non-commercial research and education use, including for instruction at the authors institution and sharing with... more

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Computational Study of Transverse Slot Injection in Supersonic Flow

by debasis Chakraborty

2022, Defence Science Journal

The knowledge of transverse sonic injection flow field is very important for the design of scramjet combustor. Three dimensional Reynolds-Averaged Navier Stokes equations alongwith turbulence models are solved to find the effect of... more

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A comparison of the numerical predictions of the supersonic combustion of Hydrogen using the S-A and SST k – ω models

by V Babu

2022, Progress in Computational Fluid Dynamics, An International Journal

In the present work, the influence of the turbulence model on the numerical predictions of supersonic combustion of hydrogen in a model combustor is investigated. Three dimensional, compressible, turbulent, reacting flow calculations... more

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Pulse Combustion Facility and Its Prliminary Application in Scramjet Research

by Tan Guan Yu

2022

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Heat transfer modes in supersonic hydrogen combustion

by Cesar Olmedo

2022

A miniature wind tunnel has been built which harnesses the power to hold supersonic flows and supersonic combustion. Experiments have been performed to test the sustainability of hydrogen combustion in supersonic Mach flows. Supersonic... more

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Numerical Investigation on the Effect of Inflow Mach Numbers on the Combustion Characteristics of a Typical Cavity-Based Supersonic Combustor

by NaMrAtA BoRdOlOI

2022, Mathematical Problems in Engineering

The air-breathing engines, commonly known as Supersonic Combustor Ramjet (SCRAMJET) engines, are one of the most prominent technologies among researchers due to their high thrust-to-weight ratio. The researchers are constantly making... more

Ficure 1: (a) Schematic diagram of the scramjet combustor (Yang et al. [36-38]). (b) Isometric view of the combustor

FicurE 2: Generation of the final mesh for the parallel cavity scramjet combustor. an ideal gas. The use of finite-rate dissipation evades the Arrhenius calculations and also reduces the computational time and cost. An implicit second-order unwinding scheme is used for spatial discretization.

The introduction of turbulence in the compressible flow equation leads to the induction of two equations, as dis- cussed in the following.

FiGure 3: (a) Comparisons of the Schlieren images and flame luminosity images with the density and temperature contours obtained from the simulations. (b) Static wall distribution along the upper wall of the cavities and comparing the experimental and simulation work witl the present simulations.

Figure 4: Schematic diagram of the DLR scramjet combustor.

FicguRE 5: Comparison of the experimental and simulation velocity profiles at streamwise locations (a) 125mm and (b) 207mm. Comparison of the experimental (c) shadow image with the contours of the (d) density and (e) pressure.

Figure 6: Grid independence test (static pressure distribution along the upper wall).

Ficure 7: Mach number contour showing the nomenclature of the shock waves and their interactions.

Figure 8: (a) Mach number contours for different incoming Mach numbers. (b) Contours plot of density for different incoming Mach numbers.

Figure 9: Contour plot of (a) temperature, (b) pressure at different incoming Mach numbers, and (c) pressure for Mach 2.52 describing the different shock waves formations and flow variables.

FicureE 10: Contour plot of H,O mass fraction at different incoming Mach numbers. 6.2.2. Mixing Efficiency. The proper mixing of the species in a combustion reaction is very essential for obtaining the desirable heat release. In this paper, ¢ =1 is used for evaluating the mixing of fuel and air. The presence of a high equivalence ratio (¢ =1) is highly affected by the mixing of fuel and air. Mixing efficiency is defined as the ratio of the hydrogen mass flux that is to be burnt to the total mass flux of hydrogen injected as described by [48] stated as follows:

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Costo di riproduzione e vantaggi comparati”. La teoria del valore di Francesco Ferrara come teoria dello scambio tra "gruppi non concorrenti

by Stefano Perri

2022

In order to reduce the cost and complexity associated with fuel injection and mixing experiments for highspeed flows, and to further enable optical access to the test section for nonintrusive diagnostics, the Enhanced Injection and Mixing... more

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Experimental Investigation of the Flow Field Characteristics under Active Flow Control

by Ibrahim Olwi

2022, International Journal of Fluid Mechanics Research

This paper describes an experimental investigation of flow field characteristics of a conceptual trapped vortex combustor (TVC). It is the first time that the novel TVC is evaluated for turboshaft engine that utilizes a single cavity to... more

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Numerical simulation of flow in a combustor with penetrable cavity wall

by Marat Goldfeld

2022

The results of numerical simulation of three configurations of a supersonic combustor are presented: without a cavity, with a solid rear wall of the cavity and a penetrable (discret) rear wall of the cavity. The calculations were carried... more

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Numerical and experimental study of supersonic combustion chamber at entrance Mach number 4. Part 1: Non-reacting flow

by Marat Goldfeld

2022, MATEC Web of Conferences

Results of joint experimental and numerical study are presented for a non-reacting 3D flow in a channel with backward-facing steps (BFS) located on top and bottom walls. Experiments were conducted in a hotshort wind tunnel at entrance... more

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Numerical investigation on cavity structure of solid-fuel scramjet combustor

by Xinyan Pei

2022, Acta Astronautica

The influences of miscellaneous combustor structures for solid fuel scramjet combustion on the performance are investigated, including a detailed interaction analysis between shocks/waves and combustion. Hydroxyl-terminated polybutadiene... more

Fig. 15. Mach number distribution along axial direction for various Dg/D,,.

Here U is the conservative variable vector. The variables used are the density p, the velocity components u, v, the total energy E, and Y; is the mass fraction of species i. F, G are inviscid flux vectors and H, R are viscous ones in x-, y-directions, respectively. S is a source vector resulting from combustion reaction. The detailed descriptions for the notions are presented as follows:

A two-equation k-@ turbulent model of shear stress transport (SST) [9,10] is employed to enclose the above governing equations. the shear stress transport (SST) k-@ turbulence model developed by Menter [9,10], which is suitable for the simulation of the supersonic flow field through the studies of Hou et al. [11] and Kumaran and Babu [12]. Eq. (3) shows the formula used to obtain the turbulent viscosity. The turbulent kinetic energy (k) and specific dissipation rate (@) are obtained by Eqs. (4) and (5), respectively. The parameter values listed below are all taken from the work of Menter [9,10]. The coefficient a* damps the turbulent viscosity, caus- ing a low-Reynolds-number correction. It is given by: The blending functions, F; and Fy, are given by:

Fig. 2. Comparison of numerical and experimental pressure distribution along the combustor [3,4].

Fig. 5. Cavity density contours (kg/m?) and streamlines, (a) no cavity, (b) L/D=3, (c) L[D=4, and (d) L/D=5.

Fig. 7. Mach number distribution along axial direction for various L/D. Fig. 6. Static temperature distribution (K) of different cavity structures, (a) no cavity, (b) L/D=3, (c) L/D=4, and (d) L/D=5

Fig. 8. The comparison of shocks and waves with and without the cavity, (a) pressure distribution (pascal) for case L/D3 and (b) enlarged view of the interaction region pressure distribution (pascal) and streamlines for the no cavity case.

Fig. 10. The combustion efficiencies for different cavities. Fig. 9. The contours of total pressure loss (a) no cavity, (b) L/D=3, (c) L[D=4, and (d) L/D=5.

Fig. 11. The specific impulse for various L/D ratios.

Fig. 12. Chemical species and temperature distributions of case L/D4 for different sections, (a) cavity section, (b) cylindrical section, and (c) diverging section.

Fig. 13. Streamlines and density contours (kg/m?) for various aft wall angles (a—d) solid fuel L/D=4, Da/D,=1.5, (a) 9=15°, (b) @=45°, (c) @=60°, an (d) @=90°.

Fig. 14. The static pressure distribution along the combustor axial for various 0.

Fig. 16. Streamlines and static pressure contours (pascal) for various offset ratio variations in case of L/D=4, 9=45°, (a) Dg/D,=1, (b) Dg/D,=1.25 (c) Dg/Dy=1.5, and (d) Dg/D,=2. The effects of various offset ratios (Dg/D,,) on the cavity flow are examined in Fig. 16. The Mach number of Dg/Dy=1 is subsonic and nearly constant in the cavity section. The flow Dg/D, of 1 achieves the condition of a thermal choke downstream the cavity section, which is induced by both heat addition and the limited initial diameter of the cylindrical section. The thermal choke is easy to be formed in the supersonic combustion, which has been also observed in the experiments on SFSCRJ of Ben-Yakar et al. [3,4]. The high back pressure caused by

Fig. 17. Total pressure loss of various Dg/Du.

Fig. 18. The combustion efficiency for various combustor structures beyond the cavity.

Fig. 20. The specific impulse for various Dg/D,. Fig. 19. The specific impulses for various aft wall angles.

The comparisons of parameters for combustor exit between experiments and simulations. Table 2 According to the distinct range of the density distribution, flow of the compression effect becomes more apparent and compression strength increases gradually with length of the cavity increasing. The strongest sections of compression all locate near the head of cavities.

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Correction: Formulation and Implementation of Inflow/Outflow Boundary Conditions to Simulate Propulsive Effects

by David Rodriguez

2022, 2018 AIAA Aerospace Sciences Meeting

Boundary conditions appropriate for simulating flow entering or exiting the computational domain to mimic propulsion effects have been implemented in an adaptive Cartesian simulation package. A robust iterative algorithm to control mass... more

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