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Design and investigation of hypersonic scramjet inlet for mach number of 10 and its performance at various flight regimes.


The term SCRAMJET (Supersonic Combustion RAMJET) is coined after developing ramjet engines. As the name implies, the ramjet engine rams through the air to compress the fluid and it has no moving parts. The static pressure rise is achieved through the ramming effect of the air; the increase in Mach number results in increase in static pressure and static temperature with the loss of stagnation pressure due to the presence of shock waves. But at high supersonic speeds, Mach number greater than 5, the static temperature of the air will be relatively high if slowed down to subsonic speeds. And the high temperature freestream air when further mixed with fuel, the fuel will be decomposed not burned [1]. So the difference between scramjet and ramjet engine is that the ramjet inlet slows down the freestream air to subsonic speeds before it reaches the combustor whereas the scramjet engine slows down the freestream air to low supersonic speeds, thereby avoiding the dissociation of molecules in the combustor. After combustion the nozzle expands the air to high supersonic speeds at the exhaust.

This study involves the comparison of scramjet inlet performances, which is specifically designed for the Mach number of 10, under different flight regimes. The design model is subjected to comparison based on the performance parameters that are required for an ideal scramjet diffuser.

The two dimensional simulation is done by using k-mega compressible turbulence model. The simulations are performed at a cruise altitude of 30 km and for the two different flight regimes. Computational Fluid Dynamics (CFD) is used to study flight simulations in both steady and un- steady flow. A time-averaged, viscous, 2 Dimensional, CFD scheme used to compute aero-thermodynamic quantities including boundary layer effects [2]. A variety of turbulent models available ranging from one to three equations transport models. Oblique shock waves, expansion waves and shock wave interactions are mainly considered. Accuracy of the solution is dependent on many parameters like size of the control volume, orientation of boundaries, discretization and its order of accuracy.

II. Scramjet Inlet Principle:

The primary purpose of an inlet (also referred to as an intake or diffuser) for any air-breathing propulsion system is to capture and compress air for processing by the remaining portions of the engine. In a conventional jet engine the inlet works in combination with a mechanical compressor to provide the proper compression for the entire engine. For vehicles flying at high supersonic (3 < M < 5) or hypersonic (M > 5) speeds, adequate compression can be achieved without a mechanical compressor. Because the airflow and compression ratio for these engines are provided entirely by the inlet, an efficient design of an inlet is crucial to the success of the engine operation [3].


Scramjet inlet incorporates oblique shock waves in various external and internal blends in order to achieve the desired combinations of compression and turning. In comparison to the external compression system, this approach can use multiple internal reflections of weaker shock waves in order to accomplish the same task with less entropy increase, but the overall length must be greater. Mixed compression systems "decouple n the engine cowl angle from the amount of compression and can result in a cowl that is parallel to the freestream flow [4].

III. Design Issues And Limits:

Some of the major issues that has to be considered while designing inlets are as follows:

* Starting and Contraction Limits

* High Temperature Effects

* Viscous Phenomena

* Boundary-Layer Separation

The problem addressed here is starting and contraction limits, which could be manipulated by varying the geometry length parameters such as ramp angle, ramp length, number of ramps, cowl deflection and contraction ratio (CR). The inlet is designed in such a way that the oblique shock waves impinges at the cowl tip thereby increasing the mass flow rate (shock-on-lip condition). The following graph indicates the selection of contraction ratio with respect to Mach Number (M).


The contraction ratio can decrease beyond the Kantrowitz limit for high flight Mach numbers because the shock structure is formed of oblique shocks, thereby generating less loss than would a normal shock assumed at the throat. For example, Smart and Trexler (2004) [5] found through experiments that their inlet remained started at M = 4.68 with a throat-to-capture contraction ratio of 0.465 whereas the Kantrowitz limit indicated 0.653. The trend increases with the Mach number as shown by the experiments collected by Van Wie (2000) [6]. Therefore it is obvious from the above statement that the area ratio (Throat area to Inlet Area) decreases as the Mach number increases.

The model design parameters are found by solving analytically using governing isentropic flow relations (1) (2) (3), oblique shock wave relations (4) (5) by following the shock-on-lip condition. Equation (5) is solved for calorically perfect gas. And the Kantrowitz limit (6) is applied to designate a contraction ratio.

[T.sub.o]/T = 1 + [gamma] - 1/2 [M.sup.2] (1)






Therefore by solving the above equations the position of cowl leading edge can be located by shock-on-lip condition.

IV. Design Specifications:

The inlet geometry is designed using CATIA V5 software using geometry creation tools. Three different inlet models varying the number of ramps in the fore body section such as three ramps, two ramps and a single ramp models are designed. In this study, cowl angle is designed with no deflection and the inlet area is assumed to be unity. The design parameters such as ramp angle ([theta]), ramp length, cowl deflection, cowl position (with respect to fore body leading edge), inlet area and throat area.




V. Mesh Generation:

Meshing is done with the help of ANSYS software, where the models are meshed using the tools that contain command buttons which allows performing operations which include creating edge meshing, face meshing and boundary conditions. For the numerical study, inlet geometry parameters such as inlet ramps angles, length and number of ramps are varied.

* Meshing is done in many methods namely edge meshing, face meshing.

* Meshed edge, faces can be copied, moved, linked or disconnected from one another.

* A rectangular domain is created around the model so as to generate the flow field.

* Sizing method can be helpful for generating fine meshes along the edges of the fore body and cowl.




A. Boundary Conditions:

Once the meshing is done, the edges are named and boundary conditions are applied to it. In order to obtain the exact simulation of the flow, the boundary is named and specified its limitation.

Separated domains was selected based on several iterations were chosen. The initialize boundary condition for all the scramjet inlet models is given been chosen.

The Boundary Conditions for the models are as follows:

VI. Results And Analysis:

ANSYS FLUENT is used for the two dimensional simulation of the flow. Computations are done for the Mach numbers of 8 and 10. Boundary conditions and the atmospheric conditions are as follows:



Pressure and Velocity Contours for three, two and single ramp at the Mach Number 10 are plotted in the figures.




The following figures represent the pressure and velocity for multiple ramp inlet models at the Mach Number 8.








The purpose of this investigation is to analyse the performance of the Mach 10 design at different flight regime. After obtaining the theoretical solution from the governing equations, the inlet is designed under Kantrowitz limit using CATIA and further meshed and simulated using ANSYS. It is evident from the results that the three ramp inlet exhibits a better performance than other inlet models or with better total pressure recovery. This study assumes that the gas is calorically perfect and the shock- boundary layer interaction is neglected for further simplicity. The Turbulence model k- omega simulates the hypersonic flow conditions, thereby capturing shocks. Thus the important parameters obtained from the computational results are investigated for better performance in terms of higher kinetic energy efficiency, lesser total pressure loss and suitable combustor entry velocity.

p                Static pressure
[rho]            Density
T                Static Temperature
po               Total Pressure
[[rho].sub.o]    Total Density
To               Total Temperature
Ath              Throat Area
[A.sub.i]        Inlet Area
[beta]           Shock Angle
[theta]           Ramp Angle


[1.] Anderson, John D., 2005. Jr.: Introduction to Flight, 5th ed., McGraw-Hill Book Company, Boston.

[2.] Shock Induced Separating Flows In Scramjet Intakes., Yufeng Yao and Daniel Rincon, School of Aerospace and Aircraft Engineering, Kingston University Roehampton Vale, Friars Avenue London, UK.

[3.] Curran, E.T., 2001. "Scramjet Engines: The First Forty Years." Journal of Propulsion and Power, 17(6): 1138-1148.

[4.] Heiser, W.H. and D.T. Pratt, Ed. 1994. Hypersonic Air Breathing Propulsion. AIAA Education Series. Washington D.C, American Institute of Aeronautics and Astronautics.

[5.] Smart, M.K. and C.A. Trexler, 2004. "Mach 4 performance of hypersonic inlet with rectangular-to-elliptical shape transition," J. Propul. Power, 20: 228-293.

[6.] Van Wie, D.M., 2001. Scramjet Inlets. Scramjet Propulsion (Progress in Astronautics and Aeronautics). P. Zarchan. 189.

(1) Ajith. A, (2) Kavitha. K, (3) Vigneshwaran. J

(1,2,3) PG Scholar, Assistant Professor, Department of Mechanical Engineering, Government College of Engineering, Salem Received 25 April 2016; Accepted 28 May 2016; Available 5 June 2016

Address For Correspondence:

Ajith. A, PG Scholar, Assistant Professor, Department of Mechanical Engineering, Government College of Engineering, Salem E-mail:
Table 1: Specifications For Three Ramp Inlet

Number of Ramps            3

Ramp angles                10[degrees], 10.5[degrees], 11[degrees]
Ramp Length (m)            3, 0.708, 0.201
Throat Area ([m.sup.2])    0.125

Table 2: Specifications For Two Ramp Inlet

Number of Ramps            2
Ramp angles                10[degrees], 14[degrees]
Ramp Length (m)            3.55, 0.63
Throat Area ([m.sup.2])    0.125

Table 3 Specifications For Single Ramp Inlet

Number of Ramps            1
Ramp angles                15[degrees]
Ramp Length (m)            3.38
Throat Area ([m.sup.2])    0.125

Table 4: Boundary Conditions For The Inlets

Inlet                      Velocity inlet

Outlet Pressure            outlet
Upper boundary             Wall
Lower boundary             Wall
Fore body                  Wall
cowl                       Wall
Fluid                      Air

Table 5: Inlet Boundary Conditions For Mach Numbers 8 And 10

Altitude                   30km
Density                    0.0184 kg/[m.sup.3]
Temperature                226.5K
Velocity                   2413.339 and 3016.74 m/s
Pressure                   1197 Pa
Turbulence viscosity       0.01
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Author:Ajith, A.; Kavitha, K.; Vigneshwaran, J.
Publication:Advances in Natural and Applied Sciences
Date:May 30, 2016
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