NUMERICAL SIMULATION OF REACTING AND NON-REACTING FLOW IN SOLID FUEL SCRAMJET ENGINE COMBUSTOR

International Journal of Mechanical and Production Engineering Research and Development (IJMPERD) ISSN (P): 2249–6890; ISSN (E): 2249–8001 Vol. 10, Issue 3, Jun 2020, 2095-2108 © TJPRC Pvt. Ltd.

NUMERICAL SIMULATION OF REACTING AND NON-REACTING FLOW IN SOLID FUEL SCRAMJET ENGINE COMBUSTOR RITUL. V. AMBETKAR1 & YASH PAL2 1 2

M.Tech Scholar, School of Aeronautical Sciences, Hindustan Institute of Technology and Science, Chennai, India

Associate Professor, Schoolof Aeronautical Sciences, Hindustan Institute of Technology and Science, Chennai, India

ABSTRACT In this study,computational analysis was carried out to study the reacting flow filed of solid-fuel scramjet combustor.

Two-dimensional, axisymmetric combustion geometry was considered for the analysis. The turbulent (k-Ɛ) model andtwo-step eddy dissipation reaction model wereused for modelling the combustion reactions. The paraffin wax was used as a solid fuel. The non-reacting and reacting flow simulation were carried out using ANSYS FLUENT software at inlet airflow of Mach number 2. The non-reacting flow analysis showedthe formation of expansion and shocks waves inthe flow. The fluctuations inflow velocities wereascribed to the formation of the shock waves. Theseshocks were the main reason for rising the pressure and temperature inside the combustor. The study of reacting flow revealed that the shock train observed due to interaction of the flow with flame and formation of combustionby-products. The high

KEYWORDS:SCRAMJET, CFD, SFSCRJ, Supersonic Combustion & k-ɛ Model

Received: Jun 05, 2020; Accepted: Jun 25, 2020; Published: Jul 03, 2020; Paper Id.: IJMPERDJUN2020197

1. INTRODUCTION

Original Article

concentrations of combustion products at the wall also suggested that the reaction takes place in the vicinity of the wall.

A scramjet combustor is the simplestand most reliable air-breathing propulsion system for supersonic flight. With the adoption of supersonic combustion, it is the only available air-breathing propulsion system tosustained hypersonic flight beyond Mach number 5. Asimple solid fuel scramjet engineconsists of an air intake system, an exhaust nozzle, and a combustor in which solid fuel is stored[1]. At the intake section, the high enthalpy air enters at supersonic speed followed by recirculation zone where it circulates near the wall. However, the flow at the centre of combustor remains supersonic. At this section,the flame is restricted using flame holders and support selfignition of the solid fuel. The combustor section allows the fuel to hold with a port diameter smaller than the inlet section, this avoids acceleration of the flow in the combustor. The final section is the divergent cone which allows the supersonic expansion (Figure 1).

Figure 1: Solid Fuel Scramjet (SFSCRJ) Engine

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The ramjet engine constitutes a prerequisite for the development of scramjet technology. It is worth mentioning that a scramjet couldhypothetically work both as a combined ramjet/scramjet engine at the same time [2]. To demonstrate fully supersonic combustion,a scramjet engine is predictable to achievea speed of about Mach 4-5.

A scramjet has no moving parts or only a few which causes the engine itself to have relatively low production costs. A difference between the hypersonic air-breathing engines and rocket engines is that the former manages to avoid the need to carry an oxidizer for fuel combustion. Elimination of oxidizer to carry on board, the vehicle would be lighter and eventually capable of carrying more payload. Unfortunately, there are also a variety of disadvantages. A scramjet cannot produce thrust unless first accelerated to higher velocities.It could though, as earlier suggested, operate as a ramjet at a lower speed. Besides, the suspension of these engines requires various structures, as well as all required control systems[3]. All the secondary equipment required to carry the vehicle to speeds suited for scramjet operation makes the entire craft heavy.

As with a rocket that travels almost vertically through the atmosphere on its way to orbit, a scramjet will take a flatter trajectory[4].Since the thrust-to-weight ratio of a scramjet engine is small compared to conventional rockets, it takes more time for the scramjet to accelerate. A very low trajectory ensures the aircraft remains in the atmosphere at hypersonic speeds for a comparatively long period. Consequently, a scramjet-propelled vehicle faces the major problems of heat resistance not just at re-entry but in its orbit trajectory. A few of the biggest process of designing a scramjet engine relates to combustion process[5]. Although many scramjet combustors can only combust fractions of the fuel supplied and produce a small amount of useful heat. Therefore, an important challenge in scramjet design is the enhancement of the combustion process. The purpose of this work is to gain more understanding of the intricate reacting flow occurring in the combustion chamber. It can be concluded altogether that enormous development work continues before scramjets are qualified for implementation in space application. Methodology A model of turbulence which is very successful and widely used is the so-called k-ɛ model.It is a two-equation model which means two additional governing equation are included to define the turbulent flow characteristics[6]. This allows the model to take into account certain effects, such as convection and turbulent diffusion of energy. The transported variables are the turbulent kinetic energy, k and its dissipation per unit time.The equations for k and ɛ are as follows: 𝜕(𝜌𝑘) 𝜕𝑡

2

𝜇

+ ∇ ∙ (𝜌𝑢𝑘) = − 3 𝜌𝑘∇. 𝑢 + 𝜎 ⊗ ∇𝑢 + ∇. [(𝑝 ) ∇𝑘] − 𝜌𝜀 𝑟𝑘

(1)

and 𝜕(𝜌𝜀) 𝜕𝑡

2

𝜇

3

𝑝𝑟𝑐

+ ∇. (𝜌𝑢𝜀) = − ( 𝑐𝜀1 − 𝑐𝜀3 ) 𝜌𝜀∇. 𝑢 + ∇. [(

𝜀

) ∇𝜀] + [𝑐𝜀1 𝜎 ⊗ ∇𝑢 − 𝑐𝑐2 𝜌𝜀] 𝑘

(2)

These are the standard k-ε equations with some extra terms. The quantities, 𝑐𝜀1 ,𝑐𝜀2 ,𝑐𝜀3 , 𝑝𝑟𝑘 and 𝑝𝑟𝑐 are experimentally determined constants[7].

The Eddy Dissipation Model (EDM) given by Magnussen and Hjertager is closely related to the Eddy Breakup Model (EBU) of Spalding [8]. It assumed that diffusion flames carry fuel and oxidizer through separate eddies. There are also rapid chemical reactions during which the fuel and the oxidizer undergo mixing at a molecular scale to perform the combustion reaction.Hence, the speed at which reactions occur depends on the degree at which turbulent eddies containing fuel and oxidizer are mixed assumingthe rapid chemical reaction is restricted in the EDM.

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2. GEOMETRY The geometry is based on Biao et al. (2013), experiments[9]. The combustor geometry is presented in Figure 2. The SFSCRJ model framework includes an isolator and a combustor. In front of the combustor, an isolator is present which has a constant area duct designed to hold the pre-combustion shock wave and to avoid contacts between the flow at the inlet and the combustor. This combustion chamber comprises of a flame-holding area near the downstream end of the area with a forward slop, a uniform cylindrical cross-sectional area and, along wildly divergent section to sustain supersonic flow without the formation of any shocks.

Figure 2: Combustor Geometry

3. BOUNDARY CONDITIONS All the simulations are carried out using ANSYS FLUENT. The Navier-Stokes equations and multi-species model are the governing equations used in this analysis. The k-ɛ and shear-stress transport (SST)turbulence modelwas used.

The wall has been fixed as an adiabatic wall without any slip boundary condition. The oxygen and nitrogen are supplied through air inlet with 0.211 and 0.789 mass fraction ratios, respectively. The pressure at the outlet segment and other flow quantities are extrapolated from the internal domain. Table 1 shows thedetailed inlet boundary conditions. The convergence of the solution is tracked and measured by different parameters of the residual flow properties, mass conservation of the computational domain, and changes in the profile of the static pressures.

Table 1: Boundary Conditions for the Combustor Inlet Property Units Inlet Pressure 10 atm Inlet Temperature 1200k Inlet Mach No. 2

4. GRID INDEPENDENT STUDY Grid refining was carried out using three different types of mesh. The first is a coarse mesh with30,000 cells, whilst the medium and fine mesh each were considered with60,000 and100,000 elements, respectively. The mesh size andnumber of elements of the modelasshown in Table.2. The flight Mach 2 was considered as intake conditions for grid refining. Results show a small difference between the medium and the fine grid solutions.Figure. 4 revealed that the further refinement of the grid is not necessary since it would not make a significantdifference in the solutions. A medium grid appears to be appropriate amongthe three grids being compared and will be used in the rest of this analysis.

Table 2: Meshing Details of Combustor Geometry Type of Mesh Mesh Sizing(mm) No. of Elements Coarse 0.25 30,000 Medium 0.2 60,000 Fine 0.15 1,00,000 www.tjprc.org

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(a)

(b)

(c) Figure 3: Mesh (a) Coarse, (b) Medium, (c) Fine

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Figure 4: Results of Static Pressure Distribution Along Wall

5. RESULTS AND DISCUSSIONS 5.1.Non-Reacting Flow 5.1.1Static pressure Contours for Different Grids The static pressure contours are presented in Figure. 5. In all the conditions, due to the convergent nozzle, a substantial increase in static pressure was observed inside the flame holding zone. The high pressure perpetuates upstream, resulting in pre-combustion shocks being formed in the isolator. The fluctuations and turbulence spread upstream through a fully subsonic field andled to a constant propagation of static pressure inside the flame retaining area.

(a)

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(b)

(c) Figure 5: Static Pressure Contour for Non-reacting flow(a) Coarse (b) Medium (c)Fine 5.2.Mach Number and Temperature Contours for Different Grids.

The non-reacting flow field was determined by setting a solid fuel asa wall boundary.Figure. 6shows the cold flow field of combustor geometry without adding fuel or its combustion. From Mach number contours, it canbe observed that at step corner the inflow air was expanded, and the number of Mach raised from 1.3 to 2.5. However, the velocity is quite low, and thus the static temperature inside the recirculation area of the combustor is comparatively high, at about 1150 K (Figure. 8). The high temperature suffices for the solid grain to ignite. The Mach number in the cylindrical section was found to be 2.4. Whereas, in the divergent zone, the flow is largely expanded withhigh velocities imparted by combustion products exiting at up to Mach 2.85.

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(a)

(b)

(c) Figure 6: Mach Number contour For Non-reacting flow(a) Coarse (b) Medium (c)Fine

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Figure 7: Mach Number at Axis for Different Grids In Figure. 7,it can be observed that the grid size has no significant effect on Mach number and at the axis of combustorthere is no direct interaction of the flow with the wall except the formation of shocks.

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(c) Figure 8: Temperature contour of combustor(a) Coarse (b) Medium (c)Fine

6. REACTING FLOW In the reacting flow analysis, the total heat flux was calculated from the radiative and convective heat transfer, whereas the conduction heatexchange within the grain wasneglected. While simulating reacting flow, pyrolysis of the solid paraffin is expected to produceC3 H8as the main gaseous product. The C3H8 combustion process is modelled witha simple two-step reaction mechanism which is given by: C3H8 + 3.5O2 → 3CO + 4H2O

(3)

CO + 0.5O2 → CO2

(4)

The EDM combustion model was used tosimulate the reacting flow. It was assumed that the reaction rates werebeing regulated by the turbulence, therefore, complex Arrhenius chemical kinetic equation was eliminated in this method. The wall wasconsideredas an adiabatic wall without any slip. The inlet boundary condition was set as a pressure inlet condition.

The air comprises of a blend of nitrogen and oxygen with 0.789 and 0.211 mass fractions, respectively were supplied at the inlet. The inflow boundary conditions werekept the same as in case of non-reacting flow. The mass flow of fuel evaporating from a solid surface was taken as 0.0197 kg/s.

6.1.Static Pressure Contour for Reacting Flow The static pressure wasfound to increase inside the flame holding area which was ascribed tothe heat liberated due to combustion of solid-fuel. High-pressure propagate in theupstream direction, resulting in the formation of pre-combustion shock inside of the isolator (Figure 9). The fluctuation in chamber pressurecan perpetuate upstream through a completely subsonic region, leading to a constant distribution of the static pressure in the flame holding area. In the case of reacting flow, the pressure was significantly increased due to the combustion reaction of fuel and oxygen species, when compared to the non-reacting case (Figure. 10).

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Figure 9: Static Pressure Contour for Reacting flow

Figure 10: Static Pressure distribution along with the wall comparison between reacting and non-reacting Flow 6.2.Mach Number & Temperature Contours for Reacting Flow Figure. 11 presents Mach number distribution of the airflow for the current reacting condition. When comparing the case with the non-reacting flow (Figure 6b), a significant change in the velocity distribution wasobserved. The Mach number increased from 2 to 3 in the non-reacting case, nearly vanishes in reacting flow analysis, leading in a peak Mach number value of 2.1. In the non-reacting case, broad areas of supersonic flow werereduced to subsonic velocities due tothe release of heat from a combustion reaction. When combustion occurs, a mixture of supersonic (near the centre) and subsonic (at the wall) stream is established near combustor exit, substituting the non-reacting supersonic flow.

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Figure 11: Mach Number Contour for Reacting Flow The temperature contour of the reacting flow is shown in Figure. 12. The results of the computation analysis revealed that the sustained combustion for inlet and geometry under investigation may exist within the combustor. Flame diffusion was observed where the temperature is 2630 K. Thisprovides a considerable amount of extra heat energy inside of the chamber. The temperature at the axis is still low, whereas high temperatures arescattered alongside the solid-fuel boundary. Thisresult shows that the combustion was taking place near the combustor wall.

Figure 12: Temperature Contour for Reacting Flow 6.3.Mass Fraction Figure. 13 and Figure. 14 shows the C5H8 and CO2 mass fractions, respectively. The C5H8 maintains high mass fraction in the recirculation zone as compared to other zones. Non-reacting fuel having an average mass fraction of 0.045. Figure 14displayed that the CO2 mass fraction spikes at 0.2 close to the exit of the combustion chamber where the most violent reaction is observed. The CO2 mass fraction distribution shows that the combustion happens again near the solid fuel wall.

A clear decrease of reactants i.e. C5H8 and O2 can be observed in Figure. 13 and Figure. 15, respectively, until theexit ofthe combustor and at the nozzle end, a rise in the concentration of combustion products such asCO2 and H2O can be seen inFigure16. www.tjprc.org

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Figure 13: C3H8 Mass Fraction Contours

Figure 14: CO2 Mass Fraction Contours

Figure 15: O2 Mass Fraction Contours

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Figure 16: H2O Mass Fraction Contours The oxygen mass fraction drops from peak reachingin the centreline to nearly zero after the flame. The reduction in the recirculation region wasquite steep and become more modest at the downstream. Also, the fuel vapours mass fraction slowly reduced from the walls towards the flames. At the exit, there wasstill a small amount of unburned fuel vapours indicating that fuel gasses are not completely burnt. The operational parameters of the close stochiometric fuel/air ratio, relatively small amount of the CO mass fraction wasobserved (peak value of 0.009), arising due to high CO conversion rate into CO2.

Figure 17: Mass fraction Distribution along the wall

7.CONCLUSIONS Computational analysis of the solid-fuel scramjet combustor operating at a velocity of Mach 2 showed that the flowwas supersonic throughout the combustor. The multiple expansion and compression waves wereformed inside the combustion chamber under supersonic conditions, triggering fluctuation of static pressure along the direction of the axis. www.tjprc.org

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When comparing the distribution of flow Mach number fornon-reacting and reacting flow, a substantial change in the velocity contourswere observed. In non-reacting flow, the inlet shock which was resulted whenthe Mach number increased from 2 to a maximum value of 2.84, nearly vanishes in reacting flow. The supersonic flow velocity in nonreacting scenario become subsonic due to increased heat generated from the combustion reaction. When combustion occurs, a combined supersonic flow (at centre) and subsonic flow (near the circumference) is established close to combustor outlet, substituting entire non-reacting supersonic stream.

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