What is supersonic air flow?

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

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



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 =

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
Physics Preference CFD
Solver Preference Fluent
Relevance 0
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
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
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
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)
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.



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; 


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); 


for i=1:431 












%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); 


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


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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What is supersonic air flow?