What is supersonic air flow?

**Supersonic speed is the speed of an object that exceeds the speed of sound (Mach 1). For objects traveling in dry air of a temperature of 20 °C (68 °F) at sea level, this speed is approximately 343.2 m/s (1,126 ft/s; 768 mph; 667.1 kn; 1,236 km/h).**

#### Let’s Look At The **Effect of Leading Edge Blunt in Shock Wave Boundary Layer Interaction in Scramjet Intake**

This paper discusses the performance enhancement of a supersonic air intake model through implementation of a leading edge blunt to the ramp section of the model. A supersonic air intake model with a sharp ramp leading edge and a supersonic flow of Mach 4.03 is initially considered for the validation of the numerical analysis with the experimental results.

Modification of the ramp leading edge is proposed as an effective method to decrease the size of the separation bubble formed due to the interaction of the shock wave and boundary layer that causes an increase in static pressure rise, decrease in stagnation pressure loss and flow distortion. The creation of an entropy layer because of the blunt leading edge and its effect on the shockwave boundary layer interaction at the isolator entry section of the model is studied in this paper. This study demonstrates the scope of overall improvement in scramjet engine performance through suitably positioned blunt ramp leading edge.

**Keywords – **

Supersonic intake, SWBLI, separation bubble, entropy layer

**1.1 Introduction **

A scramjet engine is a direct descendant of a ramjet engine. Ramjet engines have no moving parts, instead operating on compression to slow freestream supersonic air to subsonic speeds, thereby increasing temperature and pressure, and then combusting the compressed air with fuel. Lastly, a nozzle accelerates the exhaust to supersonic speeds, resulting in thrust. Due to the deceleration of the freestream air, the pressure, temperature and density of the flow entering the burner are “considerably higher than in the freestream”.

At flight Mach numbers of around Mach 6, these increases make it inefficient to continue to slow the flow to subsonic speeds. Thus, if the flow is no longer slowed to subsonic speeds, but rather only slowed to acceptable supersonic speeds, the ramjet is then termed a ‘supersonic combustion ramjet,’ resulting in the acronym *Scramjet*. Generally a Scramjet Engine starts at a hypersonic freestream Mach no. 5.00. In order to propel to those speeds, we use turbojet engines which propel to around 3.00- 4.00 Mach and from there the ramjet picks upon and starts to propel to start the scramjet engine. If we reduce the scramjet engine starting Mach number to say 3.50 or 4.00, we can eliminate one propulsion engine, i.e., ramjet engine and thus reducing weight and complexity.

Hypersonic flight in air space is currently under the spotlight of both the military and commercial sector. However the prominent issue facing hypersonic flights is the engine limitations due to the occurrence of the shock-wave/boundary layer interactions (SBLIs) in the supersonic portion of the flow inside the supersonic engine inlets. The interactions are caused by a number of oblique and normal shock waves interacting with the boundary layer of the inlet flow consequently causing boundary layer separation and unsteady flow due to severe adverse pressure gradients. Currently the problem of SBLIs is treated by the use of a boundary layer bleed system that involves the removal of a certain amount of mass of the inlet flow.

The use of this system on the other hand is accompanied by several drawbacks that are undesirable for engine performance. The purpose of this study is to investigate the use of a novel type of flow control called leading edge blunt which is able to replace the conventional bleed system and at the same time offers similar benefits in suppressing the SBLIs as well as improving the boundary layer health at the engine intake. Shock boundary layer interactions (SBLI) is a phenomenon when a shock wave meets a boundary layer and it can be found in most of the high-speed flows. The typical case of SBLI is when a generated shock wave impinges onto a surface where the boundary layer is developing.

As a result, the shock imposes a severe adverse pressure gradient towards the boundary layer causing it to eventually thicken and creates the possibility of separation. In most cases, SBLI also causes flow instability. The consequences of the phenomena are found to be detrimental especially in high speed flows. In hypersonic flows, SBLI causes intense localized heating due to high Mach numbers that can be severe enough to destroy the body of the aero-vehicle. Supersonic combustion ramjet engine is air breathing jet engine which uses vehicle’s forward motion to compress the incoming air for supersonic combustion without a rotating compressor.

In order to generate such a database, a generic, two-dimensional, planar inlet-isolator-diffuser model was designed and fabricated to replicate the lines typical of a dual-mode scramjet integrated with a hypersonic vehicle.

**1.2 Literature Review **

Scramjet Engines are a rapidly developing technology owing to many advantages such as absence of rotating parts and high specific impulse. Combustion in these engines occurs at supersonic speeds which allow greater operational efficiency at high speeds ranging from Mach 6 up to Mach 15. Unlike Ramjet engines where the flow is decelerated to subsonic speeds via a normal shock, Scramjet engines implement an oblique shock train to achieve the pressure rise required for combustion. This concept allows for a greater pressure and temperature recovery since across normal shock the loss in stagnation pressure and temperature is maximum. The basic geometry of scramjet engine is given below

*Scramjet Engine Nomenclature (Source: **Curran et al., 2001**) *

*Curran et al., 2001*

**Anderson et al. (1990) **provided basic concepts of compressible flow like isentropic flow, Rayleigh flow, Fanno flow, formation and types of shock waves etc. in his book, “Modern Compressible Flow.”

**Menter et al. (1994) **presented a two equation eddy viscosity model which utilized the original k-ω model for the inner layer of the boundary layer and switched to k-ε in the freestream flow. In the second model, he modified the eddy viscosity in the previous model which accounts for the effect of transport of principal turbulent shear stress. This new model (Shear Stress Transport – SST) was found to greatly improve the prediction of adverse pressure gradient flows, such as flows in scramjet inlets in presence of oblique shock waves.

**Emami et al. (1995) **carried out a parametric study of inlet-isolator performance of a dual mode. Scramjet Engine. To carry out the objective, a two dimensional planar inlet isolator-diffuser model was designed and fabricated. The experimental studies were conducted in the cold flow Mach 4 blowdown facility (M4BDF). The set of inlet-isolator performance data was obtained from a total of 250 geometric configurations resulting from several interchangeable and rotating cowls as well as isolator sections of different lengths.

**Holland et al. (1995) **carried out a computational and experimental study of the internal aerodynamics of a three-dimensional sidewall-compression scramjet intake at Mach 10 and studied the effect of parameters such as aerodynamic contraction ratio (CR), cowl position and Reynolds number on its performance. The experiments were performed at in the Langley 31 inch Mach 10 tunnel for three values of contraction ratios (CR=3, 5 and 9), three Reynolds numbers (Re=0.55×106 per foot, 1.14×106 per foot and 2.15×106 per foot) and three cowl positions. He concluded that upon increasing the contraction ratio and decreasing the Reynolds number, the separation induced by glancing shocks generated by the sidewall leading edges increases. He also stated that CFD results yielded good agreements with the “on design” conditions while poor agreements were obtained for highly separated flows.

**Curran et al. (2001) **provided a detailed description of scramjet propulsion in his book published by AIAA.

**1.3 Knowledge gained from the literature **

SBLIs are one of the fluid dynamics phenomena that are significant to the development of hypersonic knowledge at this age. SBLIs are prevalent in many supersonic and hypersonic applications such as supersonic and hypersonic inlets, missiles and aircraft after bodies, etc. The effects caused by SBLI are severe, therefore large portions of effort are being spent in understanding the physics of the phenomena and the methods that can be applied to overcome SBLI. The U.S. Space Shuttle program that spanned its operation from 1981 to 2011 encountered two major problems during its early stage and interestingly, both of them were related to SBLI.

The first problem was called Shuttle Flap Anomaly and it almost caused a catastrophic failure during the space shuttle’s maiden flight. A significantly larger flap deflection was needed to stabilize the shuttle than the calculated value that was determined from the ground tests. Luckily necessary control took place and the disaster was successfully avoided. Engineers came to the conclusion that the problem was caused by the failure to take into account the influence of real-gas effects on the SBLI regions since the ground tests were conducted in cold-flow facilities. The second problem was the structural failure of the leading-edge. The dynamic loads induced by the SBLI fractured the foam from the shuttle tank and impacted on the leading-edge.

###### SWBLI in mixed compression inlet

The strong dynamic loads were not predicted in the design phase and this shows that the effects of SBLI were underestimated in the early years. Another classic case is the X 15 hypersonic airplane vehicle in 1960. Due to the severe aerodynamic heating caused by shock-wave impingements on the vehicle’s body, holes were burned on its pylon surface and through the vehicle’s body. The generation of the shock train that travels internally towards the nozzle creates a number of issues. When the shock train reflects on the inlet walls, it interacts with the developing boundary layer and causes SBLI effects.

The interactions induce adverse pressure gradients that results in boundary layer separation at several locations downstream the inlet. The thickening of the boundary layer due to the separation decreases the effective throat area and could lead to inlet unstart. Another contributing factor to the inlet unstarts is the low distortion and unsteadiness which translates to large oscillating structural loads that could also result in structural fatigue. Hence the SBLIs effects in mixed compression inlet affect not only the propulsion system but also endanger the structural strength of any aero-vehicle.

**1.4 Gaps in the Literature **

There are three main areas that these problems lay in, namely Air Inlet, Combustor, and Structures and Materials. Problems within these areas vary from inlet starting problems to the inherent difficulty of the ignition of the fuel in a supersonic flow, as the possibility of failure exists anywhere from the fuel not igniting to the possibility that the ignition could take place outside of the combustor due to the extraordinary velocity of the air in the engine. Additionally, structures that can withstand the extreme temperatures experienced during hypersonic flight combined with the additional temperatures experienced during combustion are necessary. Despite the wide range of applications possible with scramjet technology, the vehicle must first be propelled to a high enough Mach number for the scramjet to start. This requires, depending on the needed application, one or two additional propulsion systems to propel the vehicle to the needed scramjet start velocity.

Current scramjet designs target the start of supersonic combustion to be between Mach 5 & 6. However, if the necessary scramjet starting Mach number is reduced, a reduction in the number of required additional propulsion systems is possible, as the gap is bridged between the maximum possible velocity of the low-speed engine(s) and the scramjet start velocity. This would have direct advantages from the resulting reduction in overall vehicle weight, the lower mass fraction required for the propulsion system (thereby resulting in more available payload weight), and fewer systems that must work in succession reliably, thereby increasing overall vehicle safety.

The focus of this project is to address this issue of reducing the starting Mach number. Unstart in Scramjet engines is characterized by the formation of a strong normal shock wave in the combustor. This shock wave propagates upstream towards the inlet and eventually reduces significantly the mass flow rate and the thrust generated by the engine. Another expected result is that unstart is more likely if the incoming stream is at a lower Mach number. For the amount of fuel burnt, parameterized by Kc (Fraction of completed combustion); although obviously high values of Kc are more likely to lead to thermal choking, few unstarted realizations with Kc< 0:85 were observed. Unstart in a scramjet engine is also characterized by the following reasons:

Firstly, since when a supersonic flow is compressed it slows down, the level of compression must be low enough (or the initial speed high enough) not to slow the gas below Mach 1. If the gas within a scramjet goes below Mach 1 the engine will “choke”, transitioning to subsonic flow in the combustion chamber. This effect is well known amongst experimenters on scramjets since the waves caused by choking are easily observable.

Additionally, the sudden increase in pressure and temperature in the engine can lead to an acceleration of the combustion, leading to the combustion chamber exploding. Secondly, the heating of the gas by combustion causes the speed of sound in the gas to increase (and the Mach number to decrease) even though the gas is still traveling at the same speed. Furthermore, forcing the speed of air flow in the combustion chamber under Mach 1 in this way is called “thermal choking”.

**1.5 Objectives of the Work **

The ultimate objective of this research is to experimentally investigate the performance of a scramjet engine for a supersonic flow of Mach 4.03 and study the effects of the blunt leading edge of the ramp on the separation induced by SWBLI. The main elements of investigation are as follows:

1. To validate the inlet isolator diffuser geometry and simulation results of scramjet using the experimental and simulation results from the previously published work.

2. To understand the flow characteristics of the ramp leading edge blunt in the SWBLI.

3. To investigate the capabilities of the ramp leading edge blunt in controlling the flow separation induced by SBLI.

4. To determine the optimal design of the leading edge blunt geometries that would lead to the best improvement of the boundary layer that encountered SBLI in Mach 4.03 flow.

1.6 Design Elements included (At Least one apart from the marked ones) Engineering Standards* Prototype and Fabrication

Design Analysis* Experimentation

Modeling and Simulation Software Development

1.7 Realistic Constraints to be addressed (At Least two to be selected) Economic Ethical

Environmental Health and Safety

Social Manufacturability

Political Sustainability

**CHAPTER –II **

**METHODOLOGY AND EXPERIMENTAL PROCEDURE **

**2.1 Methodology **

**Effect of leading edge blunt in SWBLI **

Validation of the numerical analysis

with the experimental results

Changing the geometry of the ramp

leading edge to reduce SWBLI

Effect of the leading edge blunt of

ramp on the size of the separation zone

A comparative study of leading edge blunt of

different filet dimensions on the SWBLI

**2.1.1 **Governing Equations

To understand the physics of the fluid in motion related to any engineering problem, it is important that we develop an accurate relationship among the variations of the fluid flow properties such as pressure, temperature, velocity, density etc. at discrete points in space and time. The fluid governing equations proves a theoretical solution to how these flow properties are related to each other by either integral, differential or algebraic equations. The following three fundamental laws known as the conservation laws are used to establish the governing equations of the fluid flow.

**1. Conservation of Mass **

*Rate of change of mass + net outward mass flux *= *0 *

**2. Conservation of Momentum **

*Rate of change of momentum + net outward momentum flux = sum of forces ***3. Conservation of Energy **

*Rate of work of forces + net heat flux = change in total energy *

The corresponding governing equations are continuity equation, momentum equation and energy equation.

**2.1.2 **Turbulence modeling

Moreover, turbulence is not a feature of fluids but of fluid flows and consists of many characteristics which may vary depending on the environment. Furthermore, turbulence flow is three dimensional, chaotic, stochastic and random, hence behavior of turbulence could not be precisely defined nor predicted. Lastly, turbulence causes the formation of eddies of many different length scales. Kinetic energy of the turbulent motion is contained in the large scale structures and this energy cascades from large scale structures to smaller scale structures by an inertial and essentially viscid mechanism.

This process continues, creating smaller and smaller structures that produce a hierarchy of eddies. Eventually this process creates structures that are small enough that molecular diffusion becomes important and viscous dissipation of energy finally takes place. Turbulence does not maintain itself, but depends on its environment to obtain energy. The common source of energy for turbulent fluctuations is shear in the mean flow. Turbulence occurs when the inertia forces in the fluid become significant compared to viscous forces, and is characterized by a high Reynolds Number. However, with the development of advanced mathematical codes, super computers and new generation high-speed wind tunnel testing facilities, a number of turbulence models have been developed to simulate the turbulence phenomenon under various conditions successfully, providing further understanding of the chaotic behavior.

**2.1.3 **The SST Turbulence Model

The Shear Stress Transport (SST) Model (Menter, 1993) is currently a popular turbulence model for compressible viscous flow analysis for high Reynolds numbers internal and external flows. Many practical flows occur at conditions where compressibility effects are important. For flows under adverse pressure gradients, a turbulence model to capture boundary layer separation is important. The boundary layer separation occurs at small scales, and to capture regimes with large and small scales such as separation zones and shocks, an accurate and favorable turbulence model is required. Additionally a model should be able to predict either surface heat fluxes or shear stress to obtain accurate modeling of separation flows. The equations are given below

###### Among the k-*ε *and k-*ω *models, SST is a mixture of k-*ε *and k-*ω *models. The model is being popular for supersonic and beyond flow regimes. Stress transport models use the

Reynolds stress equations provide the turbulent stresses in the mean-momentum equations. The SST model is a first order closure model. It uses formulation of the k-*ω *equation model for the inner part of the boundary layer, and gives the model the ability to directly reach the wall through the viscous sublayer while switching to the k-*ε *formulation to model the free-stream flow. This blending of the formulations gives the smooth transition that stands out from other turbulence models used for problems associated with high Reynolds numbers.

**2.2 Simulation Procedure **

The intake considered in the present study is the same as the one experimentally studied by Emami *et al. *with a few modifications to suit the computational study. In the experimental study, the model consisted of a compression ramp, fences, and cowls of different lengths, isolator, sidewalls and an expanding section downstream of the isolator which served as a diffuser during the ramjet mode of operation.

The sketch of the model used in the present numerical simulations is shown. All dimensions in the sketch are non dimensionalized with the inlet geometric throat height. The 11° inlet compression ramp simulates the fore body of a hypersonic vehicle. The compression ramp is 9.77 in. long and changes abruptly to a flat surface which marks the inlet throat (0.4 in. high and 2 in. wide) and the beginning of the isolator section. Suitable constraints were added to the CFD model to perform the analysis:

1) Density based solver was used.

2) Energy model was turned on.

3) A pressure far field type boundary condition was considered with an inlet Mach 4.03 and temperature of 69 K as mentioned in the case study.

4) An implicit formulation was considered with an AUSM flux type as this gives an exact resolution of contact and shock discontinuities. A Green-Gauss cell based Gradient Evaluation is used to evaluate inviscid flux calculations and improve computational time. A second order upwind type flow is used as the RANS model is considered and high order spatial accuracy with respect to flow characteristics are preferred.

5) A courant number of 0.8 or less was chosen for stability purposes.

6) Convergence Requirement **– **The analysis was run for up to 100000 iterations and the corresponding residuals of the governing and transport equations were recorded.

The default ANSYS convergence criterion requires that the scaled residuals defined by governing and transport equations decrease to 10-3 for all equations except the energy equations, for which the criterion is 10-6.

7) Verification **– **The analysis is refined based on the convergence and accuracy in accordance with the experimental values.

8) Post Processing – To determine the size of the separation bubble a MATLAB code was developed to calculate the skin friction coefficient.

**1. Geometry **

A schematic representation of the supersonic air intake base model considered for the present study is shown below:

###### The supersonic combustion model was originally proposed by Emami et al. during their experimental studies for a dual mode scramjet at Mach 4 at the Langley Research Center. 250 different geometric configurations of this model were tested during their experimental studies. The model chosen was the one with 9.7° convergence angle and a cowl length of 4.4 inch as it had maximum flow distortion and a maximum separation zone from the experimental results. The geometry was designed in Solidworks 2016 and imported to Ansys Design Modeler.

**Leading edge blunt**** **

A filet of radius 2.5 inch was made on the leading edge of the ramp to reduce the SWBLI by creation of an entropy layer. Figure below shows the filet of 2.5 inch radius.

**2. Meshing **

Mesh Report – Ansys 15

Object Name | Mesh |

State | Solved |

Defaults | |

Physics Preference | CFD |

Solver Preference | Fluent |

Relevance | 0 |

Sizing | |

Use Advanced Size Function | On: Proximity and Curvature |

Relevance Center | Fine |

Initial Size Seed | Active Assembly |

Smoothing | High |

Span Angle Center | Fine |

Curvature Normal Angle | Default (18.0 °) |

Num Cells Across Gap | 1 |

Min Size | Default (6.2774e-005 m) |

Proximity Min Size | Default (6.2774e-005 m) |

Max Face Size | 4.5e-003 m |

Max Size | 8.e-003 m |

Growth Rate | Default (1.20 ) |

Minimum Edge Length | 6.238e-005 m |

Inflation | |

Use Automatic Inflation | None |

Inflation Option | First Layer Thickness |

First Layer Height | 1.45e-007 m |

Maximum Layers | 20 |

Growth Rate | 1.2 |

Inflation Algorithm | Pre |

View Advanced Options | No |

Assembly Meshing | |

Method | None |

Patch Conforming Options | |

Triangle Surface Mesher | Program Controlled |

Patch Independent Options | |

Topology Checking | Yes |

Advanced | |

Number of CPUs for Parallel Part Meshing | Program Controlled |

Shape Checking | CFD |

Element Midside Nodes | Dropped |

Number of Retries | 0 |

Extra Retries For Assembly | Yes |

Rigid Body Behavior | Dimensionally Reduced |

Mesh Morphing | Disabled |

Defeaturing | |

Use Sheet Thickness for Pinch | No |

Pinch Tolerance | Default (5.6497e-005 m) |

Generate Pinch on Refresh | No |

Sheet Loop Removal | No |

Automatic Mesh Based Defeaturing | On |

Defeaturing Tolerance | Default (3.1387e-005 m) |

Statistics | |

Nodes | 314055 |

Elements | 311529 |

Mesh Metric | None |

Grid generation for the accurate prediction of complex flow structure in the mixed compression intake has been experienced to be a critical task in this study. It has been reported that a y+ value of 100 or less is essential for the accurate prediction of high speed flows (Murugan *et al. *2015). Moreover, past studies on shock-wave boundary layer interaction phenomenon (John & Kulkarni 2014) clearly showed the utmost importance of local grid refinement in accurate capture of SWBLI affected zones.

Upon considering the chances of occurrences of multiple shock wave boundary layer interaction zones both on the ramp surface and in the isolator section, boundary layer meshing has been employed near the solid surfaces. The orthogonal quality of meshing was enhanced from 0.65 before Review 1 to 0.96, where values corresponding to 0 mean a low quality of meshing. Meshing was done in such a way so as to make the region close to the boundary layer very fine and smooth. The geometry was divided into many faces and Mapped Face Meshing was done on them. Edge sizing was done on the boundary of size 0.001m to capture the SWBLI.

**CHAPTER –III **

**RESULTS AND DISCUSSION **

**1. Validation **

Meshing as on 21-01-2018 Orthogonal Quality = 0.65 Number of Nodes = 58,875

Meshing as on 03-03-2018 Orthogonal Quality = 0.96 Number of Nodes = 3,14,055

###### A graph was plot showing static pressure distribution along the cowl. The data points were extracted from experimental results of NASA Technical paper by Emami et al. and numerical simulations results from Subash et al. Both data were compared and plotted with current results in a single graph in Matlab. The percentage error obtained in the cowl graph between current work and the experimental work was 9.7.

**xy extract **software was used to extrapolate the data points from numerical and experimental

A graph was plot showing static pressure distribution along a sharp leading ramp edge. The percentage error obtained in the ramp graph between current work and the experimental work was 8.8.

An error code was written in Matlab for comparing the results of present work and the experimental results.

**2. Contours and separation zone **

Separation bubble as seen in Mach number contour. There is a significant reduction in mach number and hence the velocity of the flow due to SWBLI affecting the performance of the scramjet engine. The size of the separation zone was determined by writing a Matlab code for calculating the skin friction coefficient. A graph was plotted between the shear stress and the x direction vector which gave the length of separation bubble along x direction. A region of negative shear stress meant flow separation and creation of separation due to SWBLI.

for i=1:431

ctaw(i)= ((tawxx(i,2))^2 + (tawny(i,2))^2)^0.5;

end

for i = 1:431

ccf(i)= ctaw(i)/(0.5*0.3871*698.5*698.5);

%cfx(i)= tawxx(i,2)/(0.5*0.3871*698.5*698.5);

end

for i=1:431

x(i)=tawxx(i,1);

end

disp(tawxx(:,1))

disp(tawxx(:,2))

disp(tawny(:,2))

disp(taw(:,2))

disp(ctaw(:))

disp(cf(:,2))

disp(ccf(:))

disp(pc(:,2))

%fileID=fopen(‘table.dat’,’w’);

%fprintf(fileID,’%12.9x %12.9x %12.9x %12.9x %12.9x %12.9x %12.9s %12.9s\r\n’,’x’,’tawxx’,’tawny’,’taw’,’ctaw’,’cf’,’ccf’,’PC’); %fprintf(fileID,’%12.9f %12.9f %12.9f %12.9f %12.9f %12.9f %12.9f %12.9f\r\n’,tawxx(:,1),

tawxx(:,2),tawny(:,2),taw(:,2),ctaw,cf(:,2),ccf,pc(:,2)); %fclose(fileID);

**Result: **** **

Length of the separation bubble was determined to be 11mm in the direction of flow.

**3. Ramp Leading Edge Blunt **

Initially a leading edge blunt of 2.5 inch radius is made. After convergence there was a significant reduction in the size of the separation bubble and less flow distortion because of creation of an entropy layer preventing SWBLI.

Static pressure distribution was plotted for the blunt leading edge of the ramp along with the initial sharp leading edge in matlab. The graph showed an increase in static pressure by **17.8% **for ramp with a radius of 2.5 inch blunt leading edge.

Hence the SWBLI decreased by creation of a blunt ramp leading edge of radius 2.5 inch. This decrease in SWBLI can be attributed to the formation of an entropy layer that interferes with the SWBLI phenomena.

**4. Gantt chart**

**References**** **

1. Borovoi, V. Y., I. Egorov, A. Y. Noev, A. Skuratov and I. Struminskaya (2011). Two Dimensional interaction between an incident shock and a turbulent boundary layer in the presence of an entropy layer. *Fluid Dynamics 46*(6), 917–934.

2. Chen, F., Chen, L. and Chang, X. Three-Dimensional sidewall compression scramjet inlet – CFD simulation and experimental comparison, AIAA-03-4741, 2003. 3. Emami, S., Trexler, C.A., Auslender, A.H. and Weidner, J.P. Experimental Investigation of Inlet-Combustor Isolators for a dual mode scramjet at a mach number of 4, May 1995, NASA Technical Paper No. NASA-TP-3502. 4. Fluent 6.3 User’s Guide, Fluent Inc., 2006.

5. Goldsmith, E. L. and Seddon, J. ed. *Practical Intake Aerodynamic Design*, AIAA Inc., New York, 1993.

6. Gaitonde, D. and Shang, J. S. A numerical study of shock-on-shock viscous hypersonic flow past blunt bodies, AIAA-90-1491, 1990.

7. Holland, S. D. and Perkins, J. N. Internal shock interactions in propulsion/Airframe Integrated Three-dimensional sidewall compression scramjet inlets, AIAA-92- 3099, 1992

8. John, B., G. Sarath, V. Kulkarni and G. Natarajan (2014). Performance comparison of flux schemes for numerical simulation of high-speed inviscid flows. *Progress in Computational Fluid Dynamics, an International Journal 14*(2), 83–96.

9. B. John and P. Senthilkumar (2018) Alterations of Cowl Lip for the Improvement of Supersonic-Intake Performance, *Journal of Applied Fluid Mechanics*, Vol. 11, No. 1, pp. 31-41, 2018

10. Krishnan, l., Sandham, N. D. and Steelant, J. Shock Wave/ Boundary Layer Interactions in a Model Scramjet Intake, *AIAA J*, July 2009, **47**, (7), pp 1680-1691. 11. Riggins, D. W. The Numerical Investigation of a Dual-Mode scramjet combustor, 12. Jannaf Joint meetings, December, 1998.

13. SIVAKUMAR, R. and BABU, V. Numerical Simulations of Flow in a 3-D Supersonic Intake at High Mach Numbers, *Defense science J*, October 2006, **56**, (4), pp 465- 476.

14. Sudharsan, N. M., Jambekhar, V. A. and Babu, V. A Validation Study of OpenFoam Using the Supersonic Flow in a Mixed Compression Intake. *J Aerospace Engineering, *2010, **224, **(G6), pp 673-679.

**A. Akshay Girish Joshi **

B.Tech in Mechanical engineering, School of Mechanical Engineering, VIT University, Vellore, 632014, India

**B. Senthil Kumar P **

Assistant Professor, School of Mechanical Engineering, VIT University, Vellore, 632014, India

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