Web Site: Email: [email protected] Volume 6 ... · The results from the computational analysis of supersonic combustion using H 2 fueled scramjet combustor with diamond-shaped strut

International Journal of Application or Innovation in Engineering & Management (IJAIEM) Web Site: www.ijaiem.org Email: [email protected]

Volume 6, Issue 4, April 2017 ISSN 2319 - 4847

Volume 6, Issue 4, April 2017 Page 157

ABSTRACT Computational analysis of supersonic combustion of hydrogen using diamond-shaped injector with air inlet of 1040K is presented in this paper. The present model is based on the species transport combustion model with the standard k-ε viscous model. As the Hydrogen fuel is injected from the diamond-shaped strut injector, it is successfully used to model the turbulent reacting flow field. By providing diamond-shaped injector, expansion wave is created which cause the proper mixing between the fuels and air and it results in complete combustion as well as pollution free of combustion. The combustion efficiency in the present work is found to be 93.8%. Keywords: CFD, combustion efficiency, diamond-shaped injector, supersonic combustion

1. INTRODUCTION Scramjet and ramjet both are designed to be used for supersonic flight; however a scramjet allows the flow through the engine to remain supersonic, whereas in a ramjet the flow is slowed to subsonic levels before it enters the combustor which is the main difference between scramjet and the ramjet. The air is at sufficiently high temperature and pressure for the fuel to combust and the resulting mixture is discharged from the engine at a higher pressure. Strut injectors are located at the channel axis and directly inject the fuel into the core of the air stream which is possible without the induction of strong shock waves. Problems occur in the mixing of the reactants, flame stability and completion of the combustion within the limited combustor length which occurs due to high speed of the supersonic flow in the combustion chamber. The flow field in the scramjet combustor is highly complex which shows that when the flight speed is low, the kinetic energy of the air is not enough to be used for the optimal compression. In a supersonic combustion scramjet or ramjet, the flow is compressed and decelerated using a series of oblique shock waves. A scramjet engine is well known as hypersonic air-breathing engine in which heat release due to combustion process occurs in the supersonic flow relative to the engine. Therefore, the flow velocity throughout the scramjet remains supersonic and thereby it does not require mechanical chocking system. The scramjet engine is composed of four main sections: the inlet, isolator, combustor and exhaust which is shown in figure 1 [1].

Figure 1 Generic scramjet engine

Sources: Final Report, Department of Engineering, Australian National University

2. HISTORICAL BACKGROUND S. Roga et al. [2, 3] mentioned that there are many types of fuel injectors for scramjet engines. The fuel that is used by scramjet is usually either a liquid or a gas. The fuel and air need to be mixed about stoichiometric proportions for efficient combustion. The main problem of scramjet fuel injection is that the air flow is quite fast which shows that,

Computational Analysis of Hydrogen-Fueled Scramjet Combustor with Diamond-Shaped

Strut Injector at Mach 4

Dr. Sukanta Roga1, Dr. K.M. Pandey2

1Associate Professor, Department of Mechanical Engineering, SJB Institute of Technology, Bengaluru

2Professor, Department of Mechanical Engineering, National Institute of Technology Silchar, Assam




there is minimal time for the fuel to mix with the air and ignite to produce thrust which require about milliseconds. Hydrogen is the main fuel used for combustion. The main important aspect in designing scramjet engines is to enhance the mixing and thus reducing the combustor length. At moderate flight Mach numbers up to Mach 10, fuel injection may have a normal component into the flow from the inlet, but at higher Mach numbers, the injection must be nearly axial since the fuel momentum provides a significant portion of the engine thrust. The injector design and the flow disturbances produced by injection also should provide a region for flame holding, resulting in a stable piloting source for downstream ignition of the fuel. The injector cannot result in too several local flow disturbance, that could result in locally high wall static pressures and temperatures, leading to increased frictional losses an severe wall cooling requirements. A number of options are available for injecting fuel and enhancing the mixing of the fuel and air in high speed flows typical of those found in a scramjet combustor. Some traditional approaches for injecting fuel are: parallel injection, normal injection, transverse injection, ramp injector, strut injector, diamond shaped strut injector, wedge shaped strut injector, strut with alternating wedge injector etc. “Thrust losses in hypersonic engines–Part-1” and “Thrust losses in hypersonic engines–Part-2” is carrioud out by Riggins et al. [4,5] and they have observed that, the shock waves, incomplete mixing and viscous effects are the main factors leading to the thrust loss in supersonic combustors, though these effects aid mixing. Strut injectors offer a possibility for parallel injection without causing much blockage to the incoming stream of air and also fuel can be injected at the core of the stream. When the flight Mach number goes above the range of 3 to 6, the use of supersonic combustion allows higher specific impulse. These elevated temperatures are sufficient enough to melt down most of the known materials [6]. The increased temperatures can cause the disassociation of the combustion products, thus in-turn limiting the temperature rise and thereby reducing thermal efficiency. The near-stagnation pressure resulting from slowing down high speed supersonic flow to subsonic conditions can advance the burner inlet pressure. As inlet pressure of burner is the highest pressure within a scramjet combustor, elevating this pressure will cause the mechanical and thermal loads to increase. P.K. Tretyakov [7] worked on “The Problems of Combustion at Supersonic Flow” and it is observed that the air flow entering a combustor will remain supersonic after the optimal compression when the flight speed is higher than a certain value and that time the efficiency of the engine will decrease with a further compression. Therefore, the combustion has to take place under the supersonic flow condition. The efficiency of heat supply to the combustion chamber based on the analysis of literature data on combustion processes in a confined high-velocity and high-temperature flow for known initial parameters is considered. The process efficiency is characterized by the combustion completeness and total pressure losses.

3. MATERIALS AND METHODS 3.1 Physical Model A mathematical model comprises equations relating the dependent and the independent variables and the relevant parameters that describe some physical phenomenon. In general, a mathematical model consists of differential equations that govern the behavior of the physical system and the geometry considered in this work is the same as the one considered by M. Deepu et al. and it is shown in figure 2[8].

Figure 2 Physical model of supersonic combustor

4. COMPUTATIONAL MODEL PARAMETER Mesh generation was performed in ICEM CFD meshing software for the current model i.e., diamond-shaped strut injector. The boundary conditions are such that the air inlet and fuel inlet surfaces are defined as pressure inlets and the exhaust is defined as pressure outlet. These conditions may be more appropriate for compressible flow. In this particular model the walls of the combustor duct do not have thicknesses. The domain is completely contained by the combustor itself; therefore there is actually no heat transfer through the walls of the combustor. The supersonic combustor model considered in the present work is show in figure 2. The combustor is 0.29 m long and 0.003 m high at fuel inlet and 0.036 m at exit. Vitiated air enters through the inlet with hydrogen being injected through the diamond-shaped strut injector. The Mach number at air inlet is 4.3 and stagnation temperature and static pressure for vitiated air are 1040K and 1 bar respectively. Fuel is injected from the base which located at nozzle exit. In addition 2ddp coupled with explicit model and turbulence and finite rate chemistry are also considered.




4.1 Validation of the model

Figure 3 Experimental shadowgraph image (top) for contour plots of density (bottom)

The present model is validated by qualitative comparison of computational image (below) with an experimental shadowgraph image (top) for the cases of hydrogen injection for contour plots of density which is shown in figure 3 and this experimental analysis were done by Oevermann [10]. With inert hydrogen injection, oblique shocks are formed at the tip of the wedge that is later reflected by the upper and lower walls. At the upper and lower walls the boundary layer is affected by the reflected oblique shocks. The boundary layer on the wedge surface separates at the base and a shear layer. In some area the reflected shock waves are deflected by the hydrogen jets.

4.2 Grid Independence Study The grid independence test is accomplished on a basis of grid. The grid was then refined by adaption based on gradients of total pressure to capture the shocks. The changes in cell, faces and nodes are 109620, 164926 and 55306 respectively. The grid independent test is shown below: Grid size (Original/ Adapted / Change) Cells (36540 / 146160 / 109620) Faces (55306 / 220232 / 164926) Nodes (18766 / 74072 / 55306)

5. MATERIALS AND METHODS The results from the computational analysis of supersonic combustion using H2 fueled scramjet combustor with diamond-shaped strut injector are discussed below:

5.1 Static Pressure

Figure 4 Contours of static pressure

Figure 5 X-Y plot of static pressure

The contour of static pressure is shown in figure 4. The recompression shocks at the upper and lower wedge corners though exist are weak. At the base of the wedge the shear layers become more pronounced due to the fact that continuous ignition occurs within these shear layers. The leading edge shock wave is reflected from the top and bottom walls but the reflected shockwave from the bottom wall is stronger compared to that from the top wall. From the analysis it is observed that, after the combustion the maximum static pressure of 1655.648 kPa is produced. There is




also a subsonic recirculation zone at the base of the wedge. The fuel is injected into this subsonic recirculation zone, which helps in flame stabilization. At the lower side of the hydrogen jet there is only a compression wave but not a shock wave. The figure 5 shows the profile between the static pressure and the position of the combustion on all conditions such as air inlet, fuel inlet and pressure outlet.

5.2 Static Temperature

Figure 6 Contours of static temperature

Figure 7 X-Y plot of static temperature

The contour of static temperature of the resulting flow is shown in figure 6. It is evident from the figure 6 that the maximum temperature of 2306K is observed in the recirculation areas which are produced due to shock wave interaction and fuel jet losses concentration and the temperature is decrease slightly along the axis. Due to combustion, the recirculation region behind the injector becomes larger as compared to mixing case which acts as a flame holder for the hydrogen diffusion flame. The leading edge shock reflected off the upper and lower combustor walls makes the setting of combustion when it hits the wake in a region where large portions of the injected fuel have been mixed up with air. The shear layers at the base of the injector becomes more pronounced with combustion due to the fact that continuous ignition occurs within these shear layers. The figure 7 shows that the profile between the static temperature and the position of the combustion on all conditions such as air inlet, fuel inlet, pressure outlet and all the walls.

5.3 Mach Number

Figure 8 Contours of Mach number

Figure 9 X-Y plot of Mach number




The contours of Mach number are shown in figure 8. From the figure 8 it is evident that after the combustion the maximum Mach number of 4.67 is observed. The figure 9 shows the profile between the Mach number and the position of the combustion on all conditions such as air inlet, fuel inlet, pressure outlet etc.

5.4 Density

Figure 10 Contours of density

Figure 11 X-Y plot of density

The contours of density are shown in figure 10. From the figure 10 it is observed that the maximum density of 0.69 kg/m3 is occurred in the tip of the fuel inlet and figure 11 shows that the profile between the density and the position of the combustion on all conditions such as air inlet, fuel inlet, pressure outlet and all the walls.

5.5 Turbulence kinetic energy

Figure 12 Turbulence kinetic energy

Figure 13 X-Y plot of Turbulence kinetic energy




The figure 12 shows the turbulence kinetic energy where the maximum value of 202459.5 m2/s2 is observed after successful combustion and the figure 13 shows the X-Y plot of turbulence kinetic energy.

5.6 Mass fraction of H2

Figure 14 Mass fraction of H2

The contour of H2 Mass fraction plot for the flow field downstream of the injector is shown in the figure 14. Alternate compression and expansion took place for the jet and was not enough to disorder the flow field much in the region near to the jet outlets. But the shock wave or expansion wave reflections interfered with the upcoming jet and localized low velocity regions were produced. Though, these regions are responsible for pressure loss of the jet, certainly enhanced the mixing and reaction. Lip height plays an important role in mixing enhancement and the maximum value of 0.919 H2 is observed after successful combustion.

5.7 Mass fraction of O2

Figure 15 Mass fraction of O2

The contour of O2 Mass fraction for the flow field downstream of the injector is shown in the figure 15. Oxygen is increased in every combustion reaction in combustion applications and air provides the required oxygen. All components other than air collected together with nitrogen. In air 21% of oxygen and 79% of nitrogen are present on a molar basis. From the figure 15 it is observed that the maximum mass fraction of O2 of 0.232 is found after combustion.

5.8 Mass fraction of H2O

Figure 16 Mass fraction of H2O

The contour of water Mass fraction for the flow field downstream of the injector is shown in the figure 16. From the figure 16 is observed that water concentration is found to be maximum value of 0.142 in the shear layer formed between the two streams of flow and the low-velocity recirculation regions within the core of the upcoming jet. Typically, when dealing the chemical reaction, it’s important to remember that mass is conserved, so the mass of product is same as the mass of reactance. Even though the element exists in different the total mass of each chemical element must be same on the both side of equation. 5.9 Combustion efficiency




Figure 17 Combustion efficiency

Combustion efficiency, comb represents how much of the hydrogen is burned in a given cross section (Ax) with respect to the total injected hydrogen. The combustion efficiency is defined by Gerlinger [11] as:

2

2

( )( ) 1

inj

gas Hcomb

H

A x uY dAx

m

�

where, is the gas density, is the mass fraction of hydrogen, is the injected hydrogen mass flux and u is the velocity component normal to the cross section. The distribution of combustion efficiency along the entire length of the combustor for the diamond-shaped strut injector is shown in figure 17. The plot starts right after the trailing edge of the circular strut injector (x = 270mm) since no hydrogen is available in upstream direction. The ignition of the fuel-air mixture takes place downstream of the trailing edge of the injector. The combustion efficiency grows near the injection region where hydrogen is rapidly mixed due to the strong stream wise vorticity. For high equivalence ratios, the combustion efficiency decreases as a consequence of the decrease in the mixing efficiency due to high values of hydrogen mass flow. As the vortices travel downstream, they become weak and their ability to spread the fuel into the surrounding flow decreases. This leads to a decrease in mixing and consequently in combustion efficiency as the mixing is not sufficient.

6. MATERIALS AND METHODS The advantage of employing the complete Navier-Stokes equations extends not only the investigations that can be carried out on a wide range of flight conditions and geometries, but also in the process the location of shock wave as well as the physical characteristics of the shock layer can be exactly determined. Neglecting the presence ofbody forces, volumetric heating and the three-dimensional Navier-Stokes equations are derived as [9].

6.1 Continuity equation

(1)

where, ρ is the density and u, v, w are the velocity vectors at x, y and z directions respectively. The momentum equations in each direction are shown below:

6.2 X-momentum equation

(2)

6.3 Y-momentum equation

(3)

6.4 Z-momentum equation

(4)

6.5 Energy equation

(5)

Assuming a Newtonian fluid, the normal stress σxx, σyy, and σzz can be taken as combination of the pressure, P and the




normal viscous stress components τxx, τyy, and τzz while the remaining components are the tangential viscous stress components whereby τxy= τyx, τxz= τzx, and τyz= τzy. For the energy conservation for supersonic flows, the specific energy, E is solved instead of the usual thermal energy, H applied in sub-sonic flow problems. In three dimensions, the specific energy, E is repeated below for convenience.

(6)

Equations 1 to 6 represent the form of governing equations that are adopted for compressible flows. The solution to the above governing equations nevertheless requires additional equations to close the system.

First, the equation of state on the assumption of a perfect gas in employed that is,

where, R is the gas constant.

Second, assuming that the air is calorically perfect, the following relation holds for the internal energy.

where, Cv is the specific heat at constant volume.

Third, if the Prandtl number is assumed constant (approximately 0.71 for calorically perfect air), the thermal conductivity can be evaluated by the following:

The Sutherland’s law is typically used to evaluate viscosity, µ which is provided by

(7)

where, µ0 and T0 are reference values at standard sea level conditions.

6.6 Generalized forms of turbulence equations

(8)

(9) where,

and D = ɛ

7. Conclusions Computational analysis of diamond-shaped strut injector with K-ɛ turbulence model could expose the flow structure of progress of Hydrogen jet through the areas disturbed by the reflections of oblique shock. The k-ε turbulence model is able to predict the in those regions where the turbulence is reasonably isotropic. It is found that the maximum temperature is produced in the recirculation areas which is produced due to shock wave-expansion, wave-jet interaction and the fuel jet losses concentration. The main attention is paid to the local intensity of heat release, which ascertains together with the duct geometry, techniques for flame initiation and stabilization, injection techniques, quality of mixing the fuel with oxidizer and the gas-dynamic flow regime.

References [1] W.H. Heiser and D.T. Pratt, “Hypersonic Airbreathing Propulsion,” AIAA Educational Series, 1994. [2] S. Roga and K.M. Pandey, “Computational Analysis of Hydrogen-Fueled Scramjet Combustor Using Cavities in Tandem Flame

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[3] K.M. Pandey, S. Roga and G. Choubey, “Numerical Investigation on Hydrogen-Fueled Scramjet Combustor with Parallel Strut Fuel Injector at a Flight Mach Number of 6”, Journal of Applied Fluid Mechanics, ISSN: 1735-3572, EISSN: 1735-3645, Vol.9, No.3, pp.1215-1220, 2016, Available online at www.jafmonline.net.




[4] D.W. Riggins, C.R. McClinton and P.H. Vitt, “Thrust Losses in Hypersonic Engines–Part 1: Methodology’’, Journal of Propulsion and Power, Vol.13, No.2, 1997.

[5] D.W. Riggins, C.R. McClinton and P.H. Vitt, “Thrust Losses in Hypersonic Engines–Part 2: Applications”, Journal of Propulsion and Power, Vol.13, No.2, 1997.

[6] D. Andreadis, “Scramjets Integrate Air and Space”, 2004. [7] P.K. Tretyakov, “The Problems of Combustion at Supersonic Flow”, West-East High Speed Flow Field Conference, November

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Vol.57, No.4, pp.367-379, July 2007. [9] J. T. Guan, H. Yeoh and C. Liu, “Computational Fluid Dynamics’’, Elsevier Inc. 2008. [10] M. Oevermann, “Numerical Investigation of Turbulent Hydrogen Combustion in a Scramjet Using Flamelet Modeling”,

Aerospace Science and Technology, Vol.4, pp.463-480, 2000. [11] P. Gerlinger, P. Stoll, M. Kindler, F. Schneider and M. Aigner, “Numerical Investigation of Mixing and Combustion

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AUTHOR

Dr. Sukanta Roga received the Ph.D degree in Mechanical Engineering from National Institute of Technology Silchar, India in 2016, M.Tech in Thermal Engineering under Mechanical Department from National Institute of Technology Silchar in 2012 and B.Tech in Mechanical Engineering from National Institute of Technology Agartala, India in 2010. Currently he is working as an Associate Professor in Mechanical Engineering Department, SJB Institute of Technology, Bengaluru, India. He has published 11 International Journal Papers, 02 National Journal Papers and 09 International Conference Papers. His research interest areas are Combustion and energy, Propulsion

and power, Scramjet combustor and Gas dynamics. Email: sukanta.me42@gmailcom

Dr. K.M. Pandey received the Ph.D degree in Mechanical Engineering in 1994 from IIT Kanpur, M.Tech in Heat Power from BHUIT, Varanasi, India in 1987 and received the B.Tech degree from BHUIT Varanasi, India in 1980. He obtained. Currently he is working as Professor of the Mechanical Engineering Department, National Institute of Technology Silchar, Assam, India. He has also worked as faculty consultant in Colombo Plan Staff College, Manila, Philippines as seconded faculty from Government of India. He has published and presented more than 300 papers in International and

National Conferences and Journals. His research interest areas are the following; Combustion, High speed flows, Fuzzy logic and neural networks, Heat transfer, Internal combustion engines, Human resource management, Gas dynamics and numerical simulations in CFD. Email: [email protected].

Web Site: Email: [email protected] Volume 6 ... · The results from the computational analysis of supersonic combustion using H 2 fueled scramjet combustor with diamond-shaped strut

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