computational analysis diamond-shaped strut injector · pdf filecomputational analysis...

IJAET International Journal of Application of Engineering and Technology

Vol-2 No.-2

ISSN: 2395-3594

COMPUTATIONAL ANALYSIS DIAMOND-SHAPED STRUTINJECTOR FOR SCRAMJET COMBUSTOR AT MACH 4.3

S. Roga1 ,K.M. Pandey2 and A.P.Singh3

1Research Scholar, Department of Mechanical Engineering, [email protected], Department of Mechanical Engineering, [email protected]

NIT Silchar, Assam-788010, India3Associate Professor, Department of Mechanical Engineering, IIST, Indore [email protected]

I. INTRODUCTION

A scramjet engine is well known as hypersonic air-breathing engine in which heat release due to combustionprocess occurs in the supersonic flow relative to theengine. Therefore, the flow velocity throughout thescramjet remains supersonic and thereby it does not requiremechanical chocking system. Scramjet is designed to beused for supersonic flight; however a scramjet allows theflow through the engine to remain supersonic, whereas in aramjet the flow is slowed to subsonic levels before it entersthe combustor which is the main difference betweenscramjet and the ramjet. Strut injectors are located at thechannel axis and directly inject the fuel into the core of theair stream which is possible without the induction of strongshock waves. Problems occur in the mixing of thereactants, flame stability and completion of the combustionwithin the limited combustor length which occurs due tohigh speed of the supersonic flow in the combustionchamber. The flow field in the scramjet combustor ishighly complex which shows that when the flight speed islow, the kinetic energy of the air is not enough to be usedfor the optimal compression. In a supersonic combustionramjet or scramjet, the flow is compressed and deceleratedusing a series of oblique shock waves. The scramjet engine

is composed of four main sections: the inlet, isolator,combustor and exhaust which are shown in figure 1 [1].

Fig. 1 Generic scramjet engine

II. HISTORICAL BACKGROUND

Pandey et al. [2, 3] mentioned that there are many types offuel injectors for scramjet combustor. The fuel that is usedby scramjet is usually either a liquid or a gas. The fuel andair need to be mixed about stoichiometric proportions forefficient combustion. The main problem of scramjet fuelinjection is that, the air flow is quite fast which shows thatthere is minimal time for the fuel to mix with the air andignite to produce thrust which require about milliseconds.The main important aspect in designing scramjet engines isto enhance the mixing and thus reducing the combustorlength. At moderate flight Mach numbers up to Mach 10,fuel injection may have a normal component into the flow

ABSTRACT

Computational analysis of diamond shaped strut injector using hydrogen with air inlet of 1300K is presented inthis paper. The present model is based on the species transport combustion model with the standard k-ω viscousmodel. As the combustion of hydrogen fuel is injected from the diamond-shaped injector, it is successfully used tomodel the turbulent reacting flow field. Due to combustion, the recirculation region behind the injector becomeslarger as compared to mixing case which acts as a flame holder for the hydrogen diffusion flame. At the base of thewedge the shear layers become more pronounced due to the fact that continuous ignition occurs within these shearlayers. The air is at sufficiently high temperature and pressure for the fuel to combust and the resulting mixture isdischarged from the engine at a higher pressure. The combustion efficiency in the present work is 94.2% which iseco friendly.

Keywords: - Combustion efficiency, diamond-shaped injector, supersonic combustion.

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Roga et al. / International journal of Application of Engineering and TechnologyVol.-2 No.-2

from the inlet, but at higher Mach numbers, the injectionmust be nearly axial since the fuel momentum provides asignificant portion of the engine thrust. The injector designand the flow disturbances produced by injection alsoshould provide a region for flame holding. The injectorcannot result in too several local flow disturbance, thatcould result in locally high wall static pressures andtemperatures, leading to increased frictional losses ansevere wall cooling requirements. A number of options areavailable for injecting fuel and enhancing the mixing of thefuel and air in high speed flows typical of those found in ascramjet combustor. Some traditional approaches forinjecting fuel are: parallel injection, normal injection,transverse injection, ramp injector, strut injector, diamondshaped strut injector, wedge shaped strut injector, strutwith alternating wedge injector etc. Riggins et al. [4, 5]worked on “Thrust losses in hypersonic engines–Part-1”and “Thrust losses in hypersonic engines–Part-2” and theyhave observed that, the shock waves, incomplete mixingand viscous effects are the main factors leading to thethrust loss in supersonic combustors, though these effectsaid mixing. Strut injectors offer a possibility for parallelinjection without causing much blockage to the incomingstream of air and also fuel can be injected at the core of thestream. When the flight Mach number goes above therange of 3 to 6, the use of supersonic combustion allowshigher specific impulse. Tretyakov [6] worked on “TheProblems of Combustion at Supersonic Flow” and it hasbeen observed that, the air flow entering a combustor willremain supersonic after the optimal compression when theflight speed is higher than a certain value and that time theefficiency of the engine will decrease with a furthercompression. Therefore, the combustion has to take placeunder the supersonic flow condition. The efficiency of heatsupply to the combustion chamber based on the analysis ofliterature data on combustion processes in a confined high-velocity and high-temperature flow for known initialparameters is considered. The process efficiency ischaracterized by the combustion completeness and totalpressure losses. Cain and Walton [7] carried out “Reviewof Experiments on Ignition and Flame Holding inSupersonic Flow” and it has been observed that, the mainattention is focused at the methods by which the fuel wasignited and combustion maintained which is particularlycommon for supersonic combustion experiments and manyexamples are found in the literature of experimentsconducted with inlet temperatures much higher thanpractical in flight. Pandey and Roga found that themaximum temperature of 2096K was found in the

reciculating areas of the scramjet combustor and byproviding strut injector, expansion wave is created whichcauses the proper mixing between the fuel and air thatresults in complete combustion. Minimum amount of OHwere found after successful combustion which showsnearly complete and eco-friend combustion [8]. The inlet-combustor interaction and flow structure through ascramjet engine at a flight Mach 6 with cavity basedinjection computationally analyzed by Pandey and Roga.Fuel is injected at supersonic speed of Mach 2 through acavity based injector. These numerical simulations areaimed to study the flow structure, supersonic mixing andcombustion for cavity based injection. Cavity is of interestbecause recirculation flow in cavity would provide a stableflame holding while enhancing the rate of mixing orcombustion. The cavity effect is discussed from a viewpoint of mixing and combustion efficiency [9]. Pandey etal.[10] worked on “CFD Analysis of Supersoniccombustion using Diamond shaped Strut Injector withstandard K-Ɛ Non-premixed Turbulence model” and it hasbeen found from the work that the maximum temperatureof 1495K and maximum velocity of 3788m/s occurred inthe recirculation areas. High turbulent intensity representsa high air-fuel mixing.

III. MATERIALS AND METHODS

PHYSICAL MODEL

A mathematical model comprises equations relating thedependent and the independent variables and the relevantparameters that describe some physical phenomenon. Ingeneral, a mathematical model consists of differentialequations that govern the behavior of the physical systemand the geometry considered in this work is the same as theone considered by Deepu et al. and it is shown in figure 2[11].

Fig. 2 Physical model of supersonic combustor

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GOVERNING EQUATIONS

The advantage of employing the complete Navier-Stokesequations extends not only the investigations that can becarried out on a wide range of flight conditions andgeometries, but also in the process the location of shockwave, as well as the physical characteristics of the shocklayer, can be exactly determined. We begin by describingthe three-dimensional forms of the Navier-Stokesequations below. Neglecting the presence of body forcesand volumetric heating, the three-dimensional Navier-Stokes equations are derived as [12]:

Continuity equation+ ( ) + ( ) + ( ) = 0 (1)Where, ρ is the density and u, v, w are the velocity

vectors at x, y and z directions. The momentum equation ineach direction is shown below:

X-Momentum equation( ) + ( ) + ( ) + ( )= + + (2)

Y-Momentum equation( ) + ( ) + ( ) + ( ) += + (3)Z-Momentum equation( ) + ( ) + ( ) + ( )

= + + (4)Energy equation( ) + ( ) + ( ) + ( )+ ( + + ) + ( ) + ( )+ (5)

Assuming a Newtonian fluid, the normal stress σxx, σyy

and σzz can be taken as combination of the pressure, P andthe normal viscous stress components τxx, τyy, and τzz whilethe remaining components are the tangential viscous stress

components whereby τxy= τyx, τxz= τzx and τyz= τzy. For theenergy conservation for supersonic flows, the specificenergy, E is solved instead of the usual thermal energy, Happlied in sub-sonic flow problems. In three dimensions,the specific energy, E is repeated below for convenience.= + 12 ( + + ) (6)

Equations 1 to 6 represent the form of governingequations that are adopted for compressible flows. Thesolution to the above governing equations neverthelessrequires additional equations to close the system. First, theequation of state on the assumption of a perfect gas inemployed that is,=Where, R is the gas constant.Second, assuming that the air is calorically perfect, thefollowing relation holds for the internal energy.e = C T

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 thermalconductivity can be evaluated by the following:=The Sutherland’s law is typically used to evaluateviscosity, µ which is provided by:

= . + 120+ 120 (7)Where, µ0 and T0 are reference values at standard sea

level conditions.

Generalized forms of Turbulence Equations( ) + ( ) + ( ) + ( ) = + ++ ( = – )( ) + ( )+ ( ) + ( ) = + ++ ( = ( − )Where, = 2 + + ++ + + + +

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And, = ϵStandard (Wilcox) k-ω turbulence model( ) + = − ∗+

+ ) (8)( ) + ( )

= −+ + )+ ) (9)

REACTION MODEL

The instantaneous reaction model assumes that a singlechemical reaction occurs and proceeds instantaneously tocompletion. The reaction used for the scramjet was thehydrogen-water reaction:

2H2 + O2 → 2H2O (10)

IV. COMPUTATIONAL MODEL PARAMETER

Mesh generation was performed in ICEM CFD meshingsoftware. The current model is diamond-shaped injectorwith. The boundary conditions are such that the air inletand fuel inlet surfaces are defined as pressure inlets and theexhaust is defined as pressure outlet. These conditions maybe more appropriate for compressible flow. In thisparticular model the walls of the combustor duct do nothave thicknesses. The domain is completely contained bythe combustor itself; therefore there is actually no heattransfer through the walls of the combustor.

The model supersonic combustor considered in thepresent work is show in figure 2. The combustor is 0.29 mlong and 0.003 m high at fuel inlet and 0.036 m at exit.Vitiated air enters through the inlet with hydrogen beinginjected through the diamond-shaped injector. The Machnumber at air inlet is 4.3 and stagnation temperature andstatic pressure for Vitiated air are 1040K and 1 barrespectively. Fuel is injected from the base which located

at nozzle exit. In addition 2ddp coupled with explicitmodel and turbulence and finite rate chemistry are alsoconsidered.

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

The present model has been validated by qualitativecomparison of computational image (below) with anexperimental shadowgraph image (top) for the cases ofHydrogen injection for contour plots of density which isshown in figure 3 and this experimental analysis were doneby Oevermann [13]. With inert H2 injection, obliqueshocks are formed at the tip of the wedge that is laterreflected by the upper and lower walls. At the upper andlower walls the boundary layer is affected by the reflectedoblique shocks. The boundary layer on the wedge surfaceseparates at the base and a shear layer formed.

GRID INDEPENDENCE STUDY

The grid independence test is accomplished on a basis ofgrid. The grid was then refined by adaption based ongradients of total pressure to capture the shocks. Thechanges in cell, faces and nodes are 438480, 658712 and220232 respectively. The grid independent test is shownbelow:

Grid size (Original/ Adapted / Change)

Cells (146160 / 584640 / 438480)

Faces (220232 / 878944 / 658712)

Nodes (74072/ 294304 / 220232)

V. RESULTS AND DISCUSSIONS,,

The results from the CFD analysis for supersoniccombustion using H2 fueled scramjet combustor withdiamond-shaped fuel injector are discussed below:

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Fig. 4 Contours of static pressure

Fig. 5 X-Y plot of static pressure

The contour of static pressure is shown in figure 4. Theleading edge shock wave is reflected from the top andbottom walls but the reflected shockwave from the bottomwall is stronger compared to that from the top wall. Fromthe analysis it is observed that, after the combustion themaximum static pressure of 3141141Pa is observed. Thefuel is injected into this subsonic recirculation zone, whichhelps in flame stabilization. At the lower side of thehydrogen jet there is only a compression wave but not ashock wave. The figure 5 shows the profile between thestatic pressure and the position of the combustion on allconditions such as air inlet, fuel inlet and pressure outlet.

Fig. 6 Contours of static temperature

Fig. 7 plot of static temperature

The static temperature contour of the resulting flow isshown in figure 6. It is observed from the analysis 6 that,the maximum temperature of 2735K observed in therecirculation areas which are produced due to shock waveinteraction and fuel jet losses concentration and thetemperature is decrease slightly along the axis. The leadingedge shock reflected off the upper and lower combustorwalls makes the setting of combustion when it hits thewake in a region where large portions of the injected fuelhave been mixed up with the air. The shear layers at thebase of the injector becomes more pronounced withcombustion due to the fact that continuous ignition occurswithin these shear layers. The figure 7 shows contour ofstatic temperature without grid.

Fig. 8. Contours of Mach number

Fig. 9 X-Y plot of Mach number

The contours of Mach number are shown in figure 8. Fromthe figure 8 it is observed that, after the combustion themaximum Mach no of 4.60 is observed. The figure 9shows the profile between the Mach number and theposition of the combustion on all conditions such as airinlet, fuel inlet and pressure outlet.

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Fig. 10 Contours of density

Fig. 11 X-Y plot of density

The contours of density are shown in figure 10. From thefigure 10 it is observed that, after the combustion themaximum density of 3.93 kg/m3 is observed in the tip ofthe fuel inlet and figure 11 shows that the profile betweenthe density and the position of the combustion on allconditions such as air inlet, fuel inlet, pressure outlet andall walls.

Fig. 12 Turbulence viscosity

Fig. 13 X-Y turbulence viscosity

The figure 12 shows the turbulence viscosity where themaximum value of 0.067 kg/m-s has been observed after

successful combustion and the figure 13 shows the X-Yplot of turbulence kinetic energy.

Fig. 14 Mass fraction of H2

The contour of H2 Mass fraction plot for the flow fielddownstream of the injector is shown in the figure 14.Alternate compression and expansion took place for the jetand was not enough to disorder the flow field much in theregion near to the jet outlets. But the shock wave orexpansion wave reflections interfered with the upcomingjet and localized low velocity regions were produced.Though, these regions are responsible for pressure loss ofthe jet, certainly enhanced the mixing and reaction. Lipheight plays an important role in mixing enhancement. Themaximum H2 of 0.5 has been observed after successfulcombustion.

Fig. 15 Mass fraction of O2

The contour of O2 Mass fraction for the flow fielddownstream of the injector is shown in the figure 15.Oxygen is increased in every combustion reaction incombustion applications and air provides the requiredoxygen. All components other than air collected togetherwith nitrogen. In air 21% of oxygen and 79% of nitrogenare present on a molar basis. From the figure 15 it isobserved that, the maximum mass fraction of O2 is 1 whichis found out after combustion.

Fig. 16 Mass fraction of H2O

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The contour of water Mass fraction for the flow fielddownstream of the injector is shown in the figure 16. Fromthe figure 16 is observed that, water concentration is foundto be maximum value of 0.08 in the shear layer formedbetween the two streams of flow and the low-velocityrecirculation regions within the core of the upcoming jet.Typically, when dealing the chemical reaction, it’simportant to remember that mass is conserved, so the massof product is same as the mass of reactance. Even thoughthe element exists in different the total mass of eachchemical element must be same on the both side ofequation.

Fig. 17 Combustion efficiency

Combustion efficiency at a given x = constant section is ameasure of how much of the fuel injected upstream hasbeen consumed at that station. This is defined as:1 − m (x)m , (11)The distribution of combustion efficiency along the entirelength of the combustor for the diamond-shaped strutinjector is shown in figure 17 where the combustionefficiency is 94.2%.

VI. SUMMARY

Computational analysis of diamond-shaped injector withk-ω turbulence model could expose the flow structure ofprogress of hydrogen jet through the areas disturbed bythe reflections of oblique shock. The k-ω turbulence modelis able to predict the fluctuations in those regions where theturbulence is reasonably isotropic. It is found from CFDanalysis that the maximum temperature observed in therecirculation areas which is produced due to shockwave-expansion, wave-jet interaction and the fuel jetlosses concentration. The main attention is paid to the

local intensity of heat release, which ascertains togetherwith the duct geometry, techniques for flame initiation andstabilization, injection techniques, quality of mixing thefuel with oxidizer and the gas-dynamic flow regime.

REFERENCES

[1] Heiser WH, Pratt DT. Hypersonic Airbreathing Propulsion. AIAAEducational Series. 1994.

[2] Pandey KM, Senior Member, IACSIT and Sivasakthivel T. CFDAnalysis of Mixing and Combustion of a Hydrogen Fueled ScramjetCombustor with a Strut Injector by Using Fluent Software. IACSITInternational Journal of Engineering and Technology. Vol. 3, No. 5,2011.

[3] Pandey KM, Reddy KK SK. Numerical Simulation of WallInjection with Cavity in Supersonic Flows of Scramjet Combustion.Journal of Soft Computing and Engineering (IJSCE). Vol.2, Issue-1,March 2012, pp.142-150.

[4] Riggins DW, McClinton CR and Vitt PH. Thrust losses inhypersonic engines–Part 1: Methodology. Journal of Propulsionand Power, Vol.13, No.2, 1997.

[5] Riggins DW, McClinton CR and Vitt PH. Thrust losses inhypersonic engines–Part 2: Applications. Journal of Propulsion andPower. Vol.13 No.2, 1997.

[6] Tretyakov PK. The Problems of Combustion at Supersonic Flow.West-East High Speed Flow Field Conference. November 2007.

[7] Cain T And Walton C. Review of Experiments on Ignition andFlame Holding in Supersonic Flow. Published By The AmericaInstitute of Aeronautics And Astronautics. Rto-Tr-Avt-007-V2.

[8] Pandey KM and Roga S. CFD Analysis of Scramjet Combustor withNon-Premixed Turbulence Model using Ramp Injector.Scientific.Net, Applied Mechanics and Materials. Switzerland. Vol.555, pp.18-25, 2014.

[9] Pandey KM and Roga S. CFD Analysis of Hypersonic Combustionof H2-Fueled Scramjet Combustor with Cavity Based Fuel Injectorat Flight Mach 6. Scientific.Net, Applied Mechanics and Materials.Switzerland. Vol. 656, pp.53-63, 2014.

[10] Pandey KM, Roga S and Singh AP(2012) CFD Analysis ofSupersonic Combustion Using Diamond Shaped Strut InjectorWith standard k-ε Non-Premixed Turbulence Model. InternationalJournal of Advanced Trends in Computer Science andEngineering. Vol.1, No.1, pp.33-42.

[11] A.p. Singh, Saket b.s.pandey, Antim Rewapati, “Computationalanalysis of de laval nozzle into a suddenly expanded duct”,International Journal Of Application of Engineering andTechnology”, Oct. 2014, Vol.-1 No.-1, Pg.-30-40.

[12] Deepu M, Gokhale SS and Jayaraj S. Numerical Modeling of ScramjetCombustor. Defense Science Journal, DESIDOC. Vol.57, No.4, pp.367-379, July 2007.

[13] Tu J, Yeoh GH and Liu C. Computational Fluid Dynamics. ElsevierInc, 2008.

[14] Oevermann M. Numerical investigation of turbulent hydrogencombustion in a scramjet using flamelet modeling. AerospaceScience and Technology. Vol.4, pp.463-480, 2000.

[15] S. Pandey, pratik sharma, moin khan, “Analysis of supersonic flowsin the de -laval nozzle at 2.1 into a suddenly expanded duct atl/d=2with cavity aspect”, International Journal of Application ofEngineering and Technology, Oct. 2014, Vol.-1 No.-1, Pg.-49-53.

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