Numerical Simulation Study on Cavity-Enhanced-Supersonic ...
5th Asia-Pacific Conference on Combustion, The University of Adelaide, Adelaide, Australia 17-20 July 2005
Numerical Simulation Study on Cavity-Enhanced-Supersonic Combustion of Upstream Fuel Injection E. Jeong 1 and I.-S. Jeung 1 1
Department of Aerospace Engineering Seoul National University, Seoul, Korea
proposed scramjet engines and flow control of supersonic nozzles and jets. The presence of a cavity on an aerodynamic surface could have a large impact on the flow surrounding it. The flowfield inside a cavity is characterized by circulating flow that increases the residence time of the flow entering the cavity. Because the drag associated with flow separation is much less over a cavity than for a bluff-body, a cavity inside s combustor makes a stable flame holder with relatively little pressure drop. A rectangular cavity driven by a free shear layer provides a well defined configuration to study the flow separation and reattachment. Researchers suggested that cavity flow oscillations can actually be used to provide enhanced mixing in supersonic shear layers. The mixing was enhanced by the acoustic disturbance and the rate of the enhancement was controlled by cavity shape while the total pressure loss was negligibly small. However before implementing such techniques, one should carefully consider and evaluate any potential thrust loss and noise generation associated with the technique because of this unsteady nature of wave propagation, the flow may become unstable, and unstable combustion in the combustor can be induced. Several control methods have been proposed to suppress the oscillations in cavity. Among others, a cavity with an angled rear wall was devised to suppress the unsteady nature of the free shear layer by eliminating the generation of traveling shocks inside the cavity. Otherwise, a small disturbance produced by a mass injector located at upstream of the cavity nearly eliminated the pressure oscillation by altering the shear layers instability characteristics[1-5]. Based on these facts, the aim of the present numerical research is to investigate the flame holding and combustion enhancement for the case of Neely et al.[5].
Abstract To achieve efficient combustion within a manageable length, a successful fuel injection scheme must provide rapid mixing between the fuel and airstreams. The aim of the present numerical research is to investigate the flame holding and combustion enhancement. Additional fuel into the cavity prevents shear flow impingement on the trailing edge of the cavity. The high temperature freestream flow mixes with the cold hydrogen fuel that is injected into the cavity and raises the fuel temperature remarkably and become to start combustion. The high pressure in the cavity due to the cavity structure and combustion leads the hydrogen fuel to upstream. The shock in the cavity to be generated by the fuel injection joins together and reflects off the ceiling wall. This makes high pressure and low mach number region and makes a small recirculation in this region. This high stagnation temperature is nearly recovered in the shear layer in front of the cavity and leads to start combustion. In the downstream of the cavity, the wall pressure drops significantly. This means that the combustion phenomenon is diminished. Because fuel lumps at the trailing edge of the cavity then it spreads after the cavity so, in this region there is a strong expansion.
1 Introduction Supersonic airbreathing engines are main components of highspeed transportation vehicles. At flight speeds beyond Mach 6, air entering the combustor must be supersonic to avoid excessive dissociation of both nitrogen and oxygen gases. Consequently, the time available for fuel injection, fuel-air mixing, and combustion is very short, of the order of 1ms. To achieve efficient combustion within a manageable length, a successful fuel injection scheme must provide rapid mixing between the fuel and airstreams. While being attractive for their contribution to mixing and flame holding in scramjet combustor, strut injectors suffer from increased drag and thermal loading. Transverse injection of fuel from the side wall causes a detached normal shock to be produced upstream of the fuel jet. As a result, considerable losses in total pressure and hence cycle efficiency can occur. Angling the injectors downstream can reduce these losses, but also reduces the mixing effectiveness and flame holding ability. One of the new concepts in scramjet research combines the fuel injectors with integrated cavity. Cavities have long been an important research topic due to their interest as a fundamental fluid dynamics phenomenon, as well as simulating practical applications on aircraft. The use of cavities in supersonic flows has also been investigated in areas such as flame holding and fuel/air mixing arrangements in
2 Numerical Formulation 2.1 Governing Equations To analyze the chemically reacting supersonic viscous flow in a scramjet engine, the fully coupled form of the species conservation equations and Reynolds averaged Navier-Stokes equations is considered. As the Reynolds number in a scramjet is very high, a fully turbulent flow can be assumed. In the present study, turbulence eddy viscosity is calculated by the Menter’s SST(Shear Stress Transport) model. The SST model combines several desirable elements of existing two-equation models. The two major features of this model are a zonal weighting of model coefficients and a limitation on the growth of the eddy viscosity in rapidly strained flows. The zonal modeling uses Wilcox’s k-ω model near solid walls and the standard k-ε model(in a k-ω formulation) near boundary layer edges and in free-shear layers. This switching is achieved with
461
a blending function of the model coefficients. The governing equations for a number of N species are summarized in conservative vector form as ∂Q ∂F ∂G ∂Fv ∂G v + + = + +W ∂t ∂x ∂y ∂x ∂y
where, ρ1 ρ1 v ω1 -ρ1u1d ρ1u ρ ρ v ω ρu d -ρ u 2 2 2 2 2 2 M M M M M d ρ N ρN v ωN -ρ N u N ρN u Q = ρu F = ρu 2 +p G = ρuv Fv = τ xx W = 0 2 ρv 0 τ xy ρv +p ρuv e β 0 (e+p)v (e+p)u x µ k ∂k/∂x ρk s1 ρvk ρuk ρuω ρω ρvω s 2 µ ω ∂ω/∂x
Figure 1 Layout of Scramjet Model and the Location of the Pressure Transducers
Figure 2 Cavity Geometry and Injector Port
2.2 Numerical Methods
Inlet Mach Number Velocity Static Pressure Static Temperature Total Enthalpy
The finite volume cell-vertex scheme is used for the spatial discretization of the governing equations. The viscous terms are expressed by a central difference method and the convective terms are expressed as a difference of the numerical fluxes at the cell interface. The numerical fluxes containing artificial dissipation are formulated using Roe’s flux difference splitting(FDS) method[6]. The complete formulation of Roe’s FDS method for multi-species chemically reacting flow is based on the method of Grossman and Cinnella extended to two-dimensional curvilinear coordinates. The MUSCL scheme is used for the extrapolation of primitive variables at the cell interface[7]. In addition, the minmod limiter function is used to overcome the severe dispersion error introduced by the higher-order extrapolation and to preserve the total variation diminishing property. By applying the LU-SSOR method, governing equations can be integrated fully implicitly by the diagonal lower and upper steps with an approximate Jacobian splitting method.
4.12 2815 m/s 86.2 kPa 1190 K 6 MJ/kg
Table 1 Freestream Condition to Scramjet Combustor Mach Number Static Temperature Static Pressure Equivalent Ratio
1.0 250K 240 kPa 0.13
Table 2 Fuel Injection Condition to Scramjet Combustor
waves in the scramjet duct. This structure is an effect caused by the presence of the cavity, because simple constant area duct with sharp leading edges and no cavity would have only weak wave structure caused by boundary layer growth. In the cavity flow, the air flow over the cavity no longer remains exactly parallel with the floor so, a shear layer separates from the upstream lip then impinges closed by the trailing edge of the cavity due to the relative high Mach number (Figure 5). The high pressure at the rear face is occurred as a result of the shear layer impingement as Figure 4. The internal cavity flow is dominated by several large and
3 Results and Discussion 3.1 Experiment Discription The scramjet model used in these experiments consisted of a 500mm long rectangular duct with a constant cross-section except the cavity of 52mm x 25mm[5]. Nominally the geometry is that of a generic combustion chamber. A single, full-width, cavity was located in the duct floor at a distance of 153mm from the leading edge of the inlet. An array of pressure tappings were distributed axially along the floor centerline at the positions indicated in Figure 1. Fuel is injected axially upstream into the cavity from the cavity ramp. The experimental tests on the scramjet combustor model were carried out in the T3 facility located at the Australian National University. T3 is a large scale free-piston driven shock tunnel. Table 1 summarizes the inlet conditions to the scramjet combustor to computational simulation. These conditions were determined using the STUBE(Shock Tube). STUBE is a 1-D, non-equilibrium chemistry nozzle expansion calculation and is used to fine the pressure, temperature, Mach number, density, flow velocity along the centerline of the nozzle. Table 2 indicates the hydrogen fuel injection condition. 3.2 No fuel Injection Process Figure 3 shows the formation of a structure of oblique shock
Figure 3 Normalized Wall Pressure Distribution
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Figure 4 Pressure Contour and Streamline in the cavity
Figure 6 Wall Pressure Distribution in Non-Reacting Flow
Figure 5 Mach Contour (The dash line is sonic line)
small recirculation zones in Figure 4. These different size recirculation zones give rise to the weak flapping flow that covers the cavity so, several oblique shocks are generated. Especially the high pressure at the rear face makes strong shock and effect the downstream flow field. The shock wave develops at this point, reflects off the ceiling of the duct, then off the floor. This process of reflecting off the top and bottom walls occurs continuously along the duct and therefore the pressure fluctuation observed as Figure 3. The recirculation zones in the cavity contain high temperature and low velocity therefore, when the fuel is injected near the cavity fuel-air mixing in the supersonic combustor is enhanced.
Figure 7 Pressure Contour in Non-Reacting Flow
Figure 8 Temperature Contour in Non-Reacting Flow 3.3 Fuel Injection Process Non-Reacting Flow Figure 6 compares the wall pressure distributions for experimental and computational simulating data in nonreacting flow. There is a significant jump of static pressure in the cavity over that measured upstream in the inlet even in the absence of combustion. This pressure rise is dominantly a product of the fuel injection of additional mass into the cavity and its effect on the local flow field. Even though the fuel is injected axially upstream, the direction of fuel injection is curved by recirculation zones in the cavity and fuel fills inside cavity and also goes upward. So, oblique shock starts to generate at the leading edge of the cavity. There is a large velocity difference between upper and down side of the cavity therefore the shear layer includes relative high temperature to set up combustion. So, the cavity can play an important role in ignition. Contrary to no fuel injection in the cavity, fuel which is injected into the cavity suppresses the flow oscillations. Because additional fuel into the cavity prevents shear flow impingement on the trailing edge of the cavity. The fuel injected on the ramp functions as a stumpy body, which makes a strong shock and reflects off the top wall. The oblique shock train is also moved axially in the duct.
Figure 9 Mach number Contour in Non-Reacting Flow
Figure 10 H2 Concentration Contour in Non-Reacting Flow
freestream flow mixes with the cold hydrogen fuel that is injected into the cavity and raises the fuel temperature remarkably and become to start combustion. Therefore cavity can be the ignition source. The high pressure in the cavity due to the cavity structure and combustion leads the hydrogen fuel to upstream (Figure 14). Therefore, this fuel layer plays a role like a wedge and makes a shock wave in front of the cavity as Figure 12. The shock in the cavity that is generated by the fuel injection joins together and reflects off the ceiling wall. This makes high pressure and low mach number region and makes a small recirculation in this region (Figure 13). The high static temperature of the incoming flow in the duct is increased further as the high stagnation temperature of the total enthalpy 6MJ/kg freestream flow. This high stagnation temperature is
Reacting Flow Figure 11 shows generally higher wall pressure distribution as compared with Figure 6 because the combustion phenomena between air and hydrogen raise the pressure in the duct. The observed combustion of both experimental and computational results is the product of self-ignition. The high temperature
463
nearly recovered in the shear layer in front of the cavity and leads to start combustion. This is the reason why Figure 15 and Figure 16 have high temperature state and OH concentration distribution in the shear layer in front of the cavity. So this technique can maintain stable flame-holding through the cavity flow. In the downstream of the cavity, the wall pressure drops significantly. This means that the combustion phenomenon is
diminished. Because fuel lumps at the trailing edge of the cavity then it spreads after the cavity so, in this region there is a strong expansion. The oblique shock train is observed when the combustion occurs in the duct. But the shock train in the reacting flow moves further upstream as compared with in non-reacting flow.
4 Conclusions The air flow over the cavity no longer remains exactly parallel with the floor and the high pressure at the rear face of the cavity is occurred even though fuel is not injected. When the fuel is injected there is a significant jump of static pressure in the cavity over that measured upstream in the inlet even in the absence of combustion. Large velocity difference exists between upper and down side of the cavity therefore the shear layer includes relative high temperature to set up combustion. So, the cavity can be the ignition source. High stagnation temperature is nearly recovered in the shear layer in front of the cavity and occurs to start combustion. So this technique can maintain stable flame-holding through the cavity flow.
5 References Figure 11 Wall Pressure Distribution in Reacting Flow [1] M. R. Gruber, R. A. Baurle, T. Mathur, K.-Y. Hsu, “Fundamental Studies of Cavity-Based Flameholder Concepts for Supersonic Combustors,” Journal of Propulsion and Power, Vol.17, No.1, January-February 2001, 146 - 153 [2] Kyung Moo Kim, Seung Wook Baek and Cho Young Han, “Numerical study on supersonic combustion with cavitybased fuel injection,” International Journal of Heat and Mass Transfer, Vol. 47, Issue 2, January 2004, 271-286
Figure 12 Pressure Contour in Reacting Flow
[3] Robert C. Murray and Gregory S. Elliott, “Characteristics of the Compressible Shear Layer over a Cavity,” AIAA Journal, Vol.39, No.5, 2001, 846-856 Figure 13 Mach number Contour in Reacting Flow
[4] R. Burnes, T. P. Parr, K. J. Wilson and K. Yu, “Investigation of supersonic mixing control using cavities - Effect of fuel injection location,” AIAA-2000-3618, AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 36th, Huntsville, AL, July 16-19, 2000. [5] A. J. Neely, C. Riley, R. R. Boyce, N. R. Mudford, A. F. P. Houwing and M. R. Gruber, “Hydrocarbon and HydrogenFueled Scramjet Cavity Flameholder Performance at High Flight Mach Numbers,” AIAA-2003-6989, 12th AIAA International Space Planes and Hypersonic Systems and Technologies, Norfolk, Virginia, Dec. 15-19, 2003
Figure 14 H2 Concentration Contour in Reacting Flow
[6] Roe, P. L., “Approximation Riemann Solvers, Parameter Vectors and Difference Schemes,” Journal of Computational Physics, Vol. 43, 1981, 357-372
Figure 15 Temperature Contour in Reacting Flow
[7] Hirsch. C., Numerical Computational of Internal and External Flows, Vol.2, John Wiley & Sons, New York, 1990 [8] Li Jian-Guo, Yu Gong, Zhang Yue, Li Ying, Qian DaXing, “Experimental Studies on Self-Ignition of Hydrogen/Air Supersonic Combustion,” Journal of Propulsion and Power, Vol.13, No.4, 1997, 538-542
Figure 16 OH Concentration Contour in Reacting Flow
464
Numerical Simulation Study on Cavity-Enhanced-Supersonic Combustion of Upstream Fuel Injection E. Jeong 1 and I.-S. Jeung 1 1
Department of Aerospace Engineering Seoul National University, Seoul, Korea
proposed scramjet engines and flow control of supersonic nozzles and jets. The presence of a cavity on an aerodynamic surface could have a large impact on the flow surrounding it. The flowfield inside a cavity is characterized by circulating flow that increases the residence time of the flow entering the cavity. Because the drag associated with flow separation is much less over a cavity than for a bluff-body, a cavity inside s combustor makes a stable flame holder with relatively little pressure drop. A rectangular cavity driven by a free shear layer provides a well defined configuration to study the flow separation and reattachment. Researchers suggested that cavity flow oscillations can actually be used to provide enhanced mixing in supersonic shear layers. The mixing was enhanced by the acoustic disturbance and the rate of the enhancement was controlled by cavity shape while the total pressure loss was negligibly small. However before implementing such techniques, one should carefully consider and evaluate any potential thrust loss and noise generation associated with the technique because of this unsteady nature of wave propagation, the flow may become unstable, and unstable combustion in the combustor can be induced. Several control methods have been proposed to suppress the oscillations in cavity. Among others, a cavity with an angled rear wall was devised to suppress the unsteady nature of the free shear layer by eliminating the generation of traveling shocks inside the cavity. Otherwise, a small disturbance produced by a mass injector located at upstream of the cavity nearly eliminated the pressure oscillation by altering the shear layers instability characteristics[1-5]. Based on these facts, the aim of the present numerical research is to investigate the flame holding and combustion enhancement for the case of Neely et al.[5].
Abstract To achieve efficient combustion within a manageable length, a successful fuel injection scheme must provide rapid mixing between the fuel and airstreams. The aim of the present numerical research is to investigate the flame holding and combustion enhancement. Additional fuel into the cavity prevents shear flow impingement on the trailing edge of the cavity. The high temperature freestream flow mixes with the cold hydrogen fuel that is injected into the cavity and raises the fuel temperature remarkably and become to start combustion. The high pressure in the cavity due to the cavity structure and combustion leads the hydrogen fuel to upstream. The shock in the cavity to be generated by the fuel injection joins together and reflects off the ceiling wall. This makes high pressure and low mach number region and makes a small recirculation in this region. This high stagnation temperature is nearly recovered in the shear layer in front of the cavity and leads to start combustion. In the downstream of the cavity, the wall pressure drops significantly. This means that the combustion phenomenon is diminished. Because fuel lumps at the trailing edge of the cavity then it spreads after the cavity so, in this region there is a strong expansion.
1 Introduction Supersonic airbreathing engines are main components of highspeed transportation vehicles. At flight speeds beyond Mach 6, air entering the combustor must be supersonic to avoid excessive dissociation of both nitrogen and oxygen gases. Consequently, the time available for fuel injection, fuel-air mixing, and combustion is very short, of the order of 1ms. To achieve efficient combustion within a manageable length, a successful fuel injection scheme must provide rapid mixing between the fuel and airstreams. While being attractive for their contribution to mixing and flame holding in scramjet combustor, strut injectors suffer from increased drag and thermal loading. Transverse injection of fuel from the side wall causes a detached normal shock to be produced upstream of the fuel jet. As a result, considerable losses in total pressure and hence cycle efficiency can occur. Angling the injectors downstream can reduce these losses, but also reduces the mixing effectiveness and flame holding ability. One of the new concepts in scramjet research combines the fuel injectors with integrated cavity. Cavities have long been an important research topic due to their interest as a fundamental fluid dynamics phenomenon, as well as simulating practical applications on aircraft. The use of cavities in supersonic flows has also been investigated in areas such as flame holding and fuel/air mixing arrangements in
2 Numerical Formulation 2.1 Governing Equations To analyze the chemically reacting supersonic viscous flow in a scramjet engine, the fully coupled form of the species conservation equations and Reynolds averaged Navier-Stokes equations is considered. As the Reynolds number in a scramjet is very high, a fully turbulent flow can be assumed. In the present study, turbulence eddy viscosity is calculated by the Menter’s SST(Shear Stress Transport) model. The SST model combines several desirable elements of existing two-equation models. The two major features of this model are a zonal weighting of model coefficients and a limitation on the growth of the eddy viscosity in rapidly strained flows. The zonal modeling uses Wilcox’s k-ω model near solid walls and the standard k-ε model(in a k-ω formulation) near boundary layer edges and in free-shear layers. This switching is achieved with
461
a blending function of the model coefficients. The governing equations for a number of N species are summarized in conservative vector form as ∂Q ∂F ∂G ∂Fv ∂G v + + = + +W ∂t ∂x ∂y ∂x ∂y
where, ρ1 ρ1 v ω1 -ρ1u1d ρ1u ρ ρ v ω ρu d -ρ u 2 2 2 2 2 2 M M M M M d ρ N ρN v ωN -ρ N u N ρN u Q = ρu F = ρu 2 +p G = ρuv Fv = τ xx W = 0 2 ρv 0 τ xy ρv +p ρuv e β 0 (e+p)v (e+p)u x µ k ∂k/∂x ρk s1 ρvk ρuk ρuω ρω ρvω s 2 µ ω ∂ω/∂x
Figure 1 Layout of Scramjet Model and the Location of the Pressure Transducers
Figure 2 Cavity Geometry and Injector Port
2.2 Numerical Methods
Inlet Mach Number Velocity Static Pressure Static Temperature Total Enthalpy
The finite volume cell-vertex scheme is used for the spatial discretization of the governing equations. The viscous terms are expressed by a central difference method and the convective terms are expressed as a difference of the numerical fluxes at the cell interface. The numerical fluxes containing artificial dissipation are formulated using Roe’s flux difference splitting(FDS) method[6]. The complete formulation of Roe’s FDS method for multi-species chemically reacting flow is based on the method of Grossman and Cinnella extended to two-dimensional curvilinear coordinates. The MUSCL scheme is used for the extrapolation of primitive variables at the cell interface[7]. In addition, the minmod limiter function is used to overcome the severe dispersion error introduced by the higher-order extrapolation and to preserve the total variation diminishing property. By applying the LU-SSOR method, governing equations can be integrated fully implicitly by the diagonal lower and upper steps with an approximate Jacobian splitting method.
4.12 2815 m/s 86.2 kPa 1190 K 6 MJ/kg
Table 1 Freestream Condition to Scramjet Combustor Mach Number Static Temperature Static Pressure Equivalent Ratio
1.0 250K 240 kPa 0.13
Table 2 Fuel Injection Condition to Scramjet Combustor
waves in the scramjet duct. This structure is an effect caused by the presence of the cavity, because simple constant area duct with sharp leading edges and no cavity would have only weak wave structure caused by boundary layer growth. In the cavity flow, the air flow over the cavity no longer remains exactly parallel with the floor so, a shear layer separates from the upstream lip then impinges closed by the trailing edge of the cavity due to the relative high Mach number (Figure 5). The high pressure at the rear face is occurred as a result of the shear layer impingement as Figure 4. The internal cavity flow is dominated by several large and
3 Results and Discussion 3.1 Experiment Discription The scramjet model used in these experiments consisted of a 500mm long rectangular duct with a constant cross-section except the cavity of 52mm x 25mm[5]. Nominally the geometry is that of a generic combustion chamber. A single, full-width, cavity was located in the duct floor at a distance of 153mm from the leading edge of the inlet. An array of pressure tappings were distributed axially along the floor centerline at the positions indicated in Figure 1. Fuel is injected axially upstream into the cavity from the cavity ramp. The experimental tests on the scramjet combustor model were carried out in the T3 facility located at the Australian National University. T3 is a large scale free-piston driven shock tunnel. Table 1 summarizes the inlet conditions to the scramjet combustor to computational simulation. These conditions were determined using the STUBE(Shock Tube). STUBE is a 1-D, non-equilibrium chemistry nozzle expansion calculation and is used to fine the pressure, temperature, Mach number, density, flow velocity along the centerline of the nozzle. Table 2 indicates the hydrogen fuel injection condition. 3.2 No fuel Injection Process Figure 3 shows the formation of a structure of oblique shock
Figure 3 Normalized Wall Pressure Distribution
462
Figure 4 Pressure Contour and Streamline in the cavity
Figure 6 Wall Pressure Distribution in Non-Reacting Flow
Figure 5 Mach Contour (The dash line is sonic line)
small recirculation zones in Figure 4. These different size recirculation zones give rise to the weak flapping flow that covers the cavity so, several oblique shocks are generated. Especially the high pressure at the rear face makes strong shock and effect the downstream flow field. The shock wave develops at this point, reflects off the ceiling of the duct, then off the floor. This process of reflecting off the top and bottom walls occurs continuously along the duct and therefore the pressure fluctuation observed as Figure 3. The recirculation zones in the cavity contain high temperature and low velocity therefore, when the fuel is injected near the cavity fuel-air mixing in the supersonic combustor is enhanced.
Figure 7 Pressure Contour in Non-Reacting Flow
Figure 8 Temperature Contour in Non-Reacting Flow 3.3 Fuel Injection Process Non-Reacting Flow Figure 6 compares the wall pressure distributions for experimental and computational simulating data in nonreacting flow. There is a significant jump of static pressure in the cavity over that measured upstream in the inlet even in the absence of combustion. This pressure rise is dominantly a product of the fuel injection of additional mass into the cavity and its effect on the local flow field. Even though the fuel is injected axially upstream, the direction of fuel injection is curved by recirculation zones in the cavity and fuel fills inside cavity and also goes upward. So, oblique shock starts to generate at the leading edge of the cavity. There is a large velocity difference between upper and down side of the cavity therefore the shear layer includes relative high temperature to set up combustion. So, the cavity can play an important role in ignition. Contrary to no fuel injection in the cavity, fuel which is injected into the cavity suppresses the flow oscillations. Because additional fuel into the cavity prevents shear flow impingement on the trailing edge of the cavity. The fuel injected on the ramp functions as a stumpy body, which makes a strong shock and reflects off the top wall. The oblique shock train is also moved axially in the duct.
Figure 9 Mach number Contour in Non-Reacting Flow
Figure 10 H2 Concentration Contour in Non-Reacting Flow
freestream flow mixes with the cold hydrogen fuel that is injected into the cavity and raises the fuel temperature remarkably and become to start combustion. Therefore cavity can be the ignition source. The high pressure in the cavity due to the cavity structure and combustion leads the hydrogen fuel to upstream (Figure 14). Therefore, this fuel layer plays a role like a wedge and makes a shock wave in front of the cavity as Figure 12. The shock in the cavity that is generated by the fuel injection joins together and reflects off the ceiling wall. This makes high pressure and low mach number region and makes a small recirculation in this region (Figure 13). The high static temperature of the incoming flow in the duct is increased further as the high stagnation temperature of the total enthalpy 6MJ/kg freestream flow. This high stagnation temperature is
Reacting Flow Figure 11 shows generally higher wall pressure distribution as compared with Figure 6 because the combustion phenomena between air and hydrogen raise the pressure in the duct. The observed combustion of both experimental and computational results is the product of self-ignition. The high temperature
463
nearly recovered in the shear layer in front of the cavity and leads to start combustion. This is the reason why Figure 15 and Figure 16 have high temperature state and OH concentration distribution in the shear layer in front of the cavity. So this technique can maintain stable flame-holding through the cavity flow. In the downstream of the cavity, the wall pressure drops significantly. This means that the combustion phenomenon is
diminished. Because fuel lumps at the trailing edge of the cavity then it spreads after the cavity so, in this region there is a strong expansion. The oblique shock train is observed when the combustion occurs in the duct. But the shock train in the reacting flow moves further upstream as compared with in non-reacting flow.
4 Conclusions The air flow over the cavity no longer remains exactly parallel with the floor and the high pressure at the rear face of the cavity is occurred even though fuel is not injected. When the fuel is injected there is a significant jump of static pressure in the cavity over that measured upstream in the inlet even in the absence of combustion. Large velocity difference exists between upper and down side of the cavity therefore the shear layer includes relative high temperature to set up combustion. So, the cavity can be the ignition source. High stagnation temperature is nearly recovered in the shear layer in front of the cavity and occurs to start combustion. So this technique can maintain stable flame-holding through the cavity flow.
5 References Figure 11 Wall Pressure Distribution in Reacting Flow [1] M. R. Gruber, R. A. Baurle, T. Mathur, K.-Y. Hsu, “Fundamental Studies of Cavity-Based Flameholder Concepts for Supersonic Combustors,” Journal of Propulsion and Power, Vol.17, No.1, January-February 2001, 146 - 153 [2] Kyung Moo Kim, Seung Wook Baek and Cho Young Han, “Numerical study on supersonic combustion with cavitybased fuel injection,” International Journal of Heat and Mass Transfer, Vol. 47, Issue 2, January 2004, 271-286
Figure 12 Pressure Contour in Reacting Flow
[3] Robert C. Murray and Gregory S. Elliott, “Characteristics of the Compressible Shear Layer over a Cavity,” AIAA Journal, Vol.39, No.5, 2001, 846-856 Figure 13 Mach number Contour in Reacting Flow
[4] R. Burnes, T. P. Parr, K. J. Wilson and K. Yu, “Investigation of supersonic mixing control using cavities - Effect of fuel injection location,” AIAA-2000-3618, AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 36th, Huntsville, AL, July 16-19, 2000. [5] A. J. Neely, C. Riley, R. R. Boyce, N. R. Mudford, A. F. P. Houwing and M. R. Gruber, “Hydrocarbon and HydrogenFueled Scramjet Cavity Flameholder Performance at High Flight Mach Numbers,” AIAA-2003-6989, 12th AIAA International Space Planes and Hypersonic Systems and Technologies, Norfolk, Virginia, Dec. 15-19, 2003
Figure 14 H2 Concentration Contour in Reacting Flow
[6] Roe, P. L., “Approximation Riemann Solvers, Parameter Vectors and Difference Schemes,” Journal of Computational Physics, Vol. 43, 1981, 357-372
Figure 15 Temperature Contour in Reacting Flow
[7] Hirsch. C., Numerical Computational of Internal and External Flows, Vol.2, John Wiley & Sons, New York, 1990 [8] Li Jian-Guo, Yu Gong, Zhang Yue, Li Ying, Qian DaXing, “Experimental Studies on Self-Ignition of Hydrogen/Air Supersonic Combustion,” Journal of Propulsion and Power, Vol.13, No.4, 1997, 538-542
Figure 16 OH Concentration Contour in Reacting Flow
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