EFFECT OF COWL ANGLE IN THE PERFORMANCE OF SCRAMJET AIR INTAKES

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1 International Journal of Mechanical Engineering and Technology (IJMET) Volume 8, Issue 11, November 2017, pp , Article ID: IJMET_08_11_091 Available online at ISSN Print: and ISSN Online: IAEME Publication Scopus Indexed EFFECT OF COWL ANGLE IN THE PERFORMANCE OF SCRAMJET AIR INTAKES Senthilkumar P, Sharan Chandran M, Sreeja Sadasivan and Govindha rasu N* School of Mechanical Engineering K Senthil kumar Mechanical and Industrial Engineering, Higher college of Technology, Muscat ABSTRACT A Scramjet engine (Supersonic Combusting Ramjet) entails combustion at supersonic flow by generating high pressure through recompression (oblique) shock train. Shock Boundary Layer Interaction (SBLI) is an undesirable phenomenon that occurs inside the scramjet intake which has a detrimental effect on its performance. Due to adverse pressure gradient created by SBLI, the flow might be decelerated to subsonic speeds via normal shock. This leads to unstarting of the engine including other undesirable phenomena such as high drag and increased localized heating. The effect of cowl angle is studied on the performance of the scramjet engine in this study. A scramjet inlet prototype is designed in Solid Works followed by CFD analysis using ANSYS tool. The effects of cowl angle, static pressure rise, total pressure recovery and Flow Distortion are brought out. Keywords: Scramjet intake, SBLI, cowl, ramp and stagnation pressure loss. Cite this Article: Senthilkumar P, Sharan Chandran M, Sreeja Sadasivan and Govindha rasu N and K Senthil kumar, Effect of Cowl Angle in the Performance of Scramjet Air Intakes, International Journal of Mechanical Engineering and Technology 8(11), 2017, pp INTRODUCTION A ramjet engine uses the kinetic energy of air to achieve compression via normal shock wave. But this pressure rise is associated with a huge stagnation pressure loss and reduction to subsonic speeds. Also, the high temperature rise creates additional losses due to dissociation of molecules. A scramjet engine is a variant of a ramjet air breathing jet engine in which combustion takes place in supersonic airflow. The scramjet is composed of three basic components: a converging inlet, where incoming air is compressed; a combustor, where gaseous fuel is burned with atmospheric oxygen to produce heat; and a diverging nozzle, where the heated air is accelerated to produce thrust. Shock wave- boundary layer interaction (SBLI) is the problem encountered by the air- breathing propulsion system. The SBLIs cause editor@iaeme.com

2 Senthilkumar P, Sharan Chandran M, Sreeja Sadasivan Govindha rasu N and K Senthil kumar boundary layer separation due to the adverse pressure gradient imposed by the shock. These result in total pressure loss, flow unsteadiness, increased aerodynamic drag and heat fluxes, fluctuating pressure loads and flow distortion in the intake section leading to the reduction of the overall propulsive efficiency of the hypersonic vehicle and premature structural fatigue of aero-structures. The quality of the flow field is also reduced significantly. Flow control mechanisms have been developed to prevent the boundary layer separation and to keep the overall propulsive efficiency high even at off-design conditions. 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-isolatordiffuser 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=0.55x10 6 per foot, 1.14x10 6 per foot and 2.15x10 6 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. Henckels et al. (2004) carried out studies on flow phenomenon on scramjet inlets wherein the influence of different inlet geometries on the heat load was determined. The test was carried out using a hypersonic blowdown tunnel H2K. Also, influence of sidewalls on the flow field was studied. Introduction of a 3D sidewall leads to a shock that glances along the ramp surface and impinges on the boundary layer of the adjacent wall. Also it creates a complex pattern of shock -shock and shock-boundary layer interactions which are difficult to predict using commercial available CFD codes. Hence for simplicity, the sidewalls are not included in the geometry used in this work. Hongjun et al. (2005) provided a method for designing a two dimensional mixed compression two ramp supersonic inlet to maximize pressure recovery and matching the inlet mass flow demand. An optimization criterion is defined for on-design conditions to maximize pressure recovery. For off design conditions, the pressure recovery depends upon the inlet configuration. A conservative value of total pressure recovery is calculated to determine the performance of the engine. The mass flow demand is matched by adjusting the second ramp angle and the bypass door. Sivakumar et al. (2006) performed numerical simulations for a 3D inlet section using commercial coding package ANSYS Fluent. The SST k-ω turbulence model was used with y+ value of 105. The results were validated by simulating flow through a 2D mixed compression hypersonic inlet as well as comparing with experimental data. From the characteristic curves obtained, the value critical angle of attack was determined to be 8⁰ with higher values of flow distortion for other values of angle of attack. Krause et al. (2006) conducted an experimental and numerical analysis on a scramjet inlet using Reynolds Averaged Flow Solvers. Two intake models were studied, one with sharp leading edge and the other with a blunt leading edge. Also, various turbulence models were investigated. The Shear Stress Transport (SST) and Spalart-Allmaras (SA) turbulence models yielded similar results while the original k-ω turbulence showed much greater separation and thicker boundary layers. The results were not validated with experimental data and only one result was presented. The result concluded that the side wall shock bends the ramp shock upstream resulting in a separation bubble which editor@iaeme.com

3 Effect of Cowl Angle in the Performance of Scramjet Air Intakes blocks the flow in the isolator. Subash et al. (2011) carried out numerical simulations on the flow through inlet and isolator of a dual mode scramjet engine at Mach 4 who used the two equation SST k-ω turbulence model in ANSYS Fluent. The SST k-ω turbulence model captures boundary layer separation under adverse pressure gradient effectively and is recommended for problems involving compressible flow. The results obtained were validated with experimental data for three different cowl lengths and four different cowl convergence angles. Soumyajit et al. (2012) carried out numerical simulations using three dimensional Reynolds Averaged Navier Stokes (RANS) solvers and the SST turbulence model. The steady state simulations were carried out for different Mach numbers ranging from 3 to 8. Effect of wall boundary conditions both isothermal (T=300K) and adiabatic on the starting Mach number was investigated. It was found that the Mach number for which the inlet unstarts was higher in case of adiabatic wall. Babinsky et al. (2011) investigated the efficiency of different vortex generators geometries in presence of a normal shock via oil flow visualisations and determination of separation regions for different upstream positions. He concluded that micro ramps do not entirely eliminate boundary layer separation but reduce it to a certain extent. Also, the corner vortices were found to be enlarged for all cases with boundary layer control the reason for which is yet to be determined. Boundary layer separation caused due to SBLI hampers the pressure recovery in the isolator and in case of high combustor back pressure, leads to unstarting of the inlet. In his book, Shock wave Boundary layer interactions he has explained the phenomenon of SBLI and its flow physics in great detail. Saad et al. (2012) carried out experimental to test the effect micro ramps on supressing the shock boundary layer interactions (SBLI). The freestream Mach number was maintained at 5 and three different configurations of micro ramps MR40, MR60 and MR80 were tested. Detailed flow field behind the micro ramp was investigated using Schlieren imaging and Infrared thermography. The micro ramp configuration MR80 was found mitigate SBLI better than the former two. Sun et al. (2012) carried out an experimental study to investigate the three dimensional flow field behind a single micro ramp in a Mach 2 supersonic flow. Manoviraj singh et al. (2017) studied the performance of Cesna-172 aircraft. Senthilkumar and Sreeja (2016) reviewed on high speed flow with micro ramps. Senthilkumar and kamalraj (2016) studied the numerical simulation of a scramjet intake with micro-cavity. A pair of streamwise vortices originating from the trailing edge of the ramp enveloped by an arc shaped external shear layer was captured in the time averaged organization. On instantaneous analysis, the enveloping vortices were found to be the Kelvin-Helmholtz (K-H) instability that bounded the wake region. A net lift force is experienced by both type of vortices. 2. COMPUTATIONAL MODELLING AND MESHING 2.1. Scramjet Geometry The geometry of the scramjet engine was designed in Solid Works. Initially, the analysis was done only for a scramjet engine alone without any blockages for cowl angles 0deg, 4deg, 5deg and 6deg respectively. The computational data obtained for the case of 0 deg was validated with the experimental values. The data included pressure values at different stream wise locations and the shock pattern obtained in the inlet-isolator. The dimensions used and the model are given below in Fig 1 and 2 respectively editor@iaeme.com

4 Senthilkumar P, Sharan Chandran M, Sreeja Sadasivan Govindha rasu N and K Senthil kumar Figure 1 Scramjet Geometry Dimensions Figure 2 Scramjet Geometry Figure 3 Dimensions of Computational Domain In the above Fig 3, the computational domain is shown with the dimensions and nomenclature of different surfaces. The spanwise width of the geometry is 30mm (origin at z=0mm) and is axisymmetric about the plane passing through z=15mm. The viewing plane in the above figure is at z=30mm which was named as side. For simplification of geometry, computational time and the resulting three dimensional flow-field, sidewalls are not included in the analysis. Instead, the surface side at z=30mm is considered as a wall with a symmetrical plane at z=15mm editor@iaeme.com

5 Effect of Cowl Angle in the Performance of Scramjet Air Intakes 2.2. MESHING A structured grid was used for all the cases simulated in ANSYS Fluent The meshing was carried out in ICEM CFD as shown in Fig 4. To capture the shock boundary layer interactions, inflation layers and appropriate y+ (the first minimum grid point from the wall) distance was maintained. For a Reynolds number of 4.8x105, estimated wall distance was 2.9x10-7 to keep the value of y+ less than 1. These values were calculated using y-plus estimation calculators available. online. The cell number varied from 60,000 for cases without blockages and micro ramps to 2.1 million for cases with blockages and micro ramps. In each case, the pre-mesh quality (in Fig 5) was obtained above 0.7 with aspect ratio less than Figure 4 Meshing with inflation layers Figure 5 Pre-mesh quality editor@iaeme.com

6 Senthilkumar P, Sharan Chandran M, Sreeja Sadasivan Govindha rasu N and K Senthil kumar 2.3. GRID INDEPENDENCY TEST Grid Independency test (GIT) was carried out on the model (scramjet cowl angle 0 deg) with both structured and unstructured mesh having same amount of cells 0.9 million. The details of each mesh are provided in the table below. The variation in the results was found to be within ±8%. The meshing images for cases with ramps and blockages including the mesh report are mentioned in the Appendix. Fig 6 shows the velocity distribution along the midsection of the isolator and Fig 7 depicts the percentage error between the velocity values of hex mesh and test mesh. Figure 6 Velocity distribution along the Isolator Figure 7 Percentage error between Hex and Tet mesh 2.4. SETUP IN ANSYS FLUENT Since the flow is compressible in a supersonic flow, a density based solver is used. The energy equation is enabled and the k-ω SST turbulence model is used to capture boundary layer separation accurately under adverse pressure gradient. This model has been widely used in numerical simulations of supersonic flows owing to the above mentioned advantage. The operating pressure is kept as 0 Pa (due to presence of vacuum inside the wind tunnel) so that the absolute pressure equals the gauge pressure. The boundary conditions are as follows: editor@iaeme.com

7 Effect of Cowl Angle in the Performance of Scramjet Air Intakes Table 1 Boundary Conditions used in the setup The outlet pressure cannot be predicted at the beginning of the calculation as the flow is supersonic i.e. the conditions at the outlet cannot be communicated upstream (Ref Sec 2.1 and 3.1). Fluent 15.0 calculates the outlet pressure gradually from the inlet conditions as the calculation proceeds. The turbulence intensity was calculated by the following empirical relation, = 0.16() 1 8 (6.1) Solution method used is Implicit AUSM solver. Discretization of Second order is used for increased accuracy. The Courant number was kept at 0.5 and the under relaxation factors as well as the discretization order were reduced initially. After a stable solution was obtained (after ~2000 iterations) they were increased gradually. All the residuals were maintained at 1e-04 and to detect convergence, a surface monitor of mass flow rate was enabled at the outlet. To capture the shock waves effectively, Adaptive grid was used using the pressuregradient. 3. RESULTS AND DISCUSSIONS 3.1. Cowl Angle 0 deg The pressure contours for cowl angle 0 deg is given in Figure 8. Figure 8 Pressure contours for cowl angle 0 deg at symmetry plane editor@iaeme.com

8 Senthilkumar P, Sharan Chandran M, Sreeja Sadasivan Govindha rasu N and K Senthil kumar As seen from Fig 8, a clean shock pattern is observed when the cowl convergence angle is 0 deg. The recompression shock train in the isolator region is also observed distinctly. The cowl is designed to produce a two-shock system with the second shock impinging just upstream of the ramp shoulder. The majority of the pressure rise is seen to be occurring from the first two shocks while remaining fairly constant in the isolator region. Also, high pressure region at the shock impingement location is observed due to shock-shock interactions between the incident and reflecting shock which also leads to high localised temperature. Figure 9 Velocity contours for cowl angle 0 deg at symmetry plane Figure 10 Shock Induced Boundary Layer Separation at symmetry plane From figs 9 and 10, shock induced boundary layer separation region is observed on the ramp surface near the shoulder region with some boundary layer thickening in the isolator section. Also, due to boundary layer separation, a separation and a reattachment shock is observed as well. In the isolator region as mentioned earlier, minor boundary layer thickening due to the pressure gradient of the oblique shock wave train is witnessed. Hence one can conclude that successive reflections lead to decrease in the shock strength and the pressure gradient across them editor@iaeme.com

9 Effect of Cowl Angle in the Performance of Scramjet Air Intakes Figure 11 Pressure distribution at ramp and cowl surface Figure 11 shows the pressure distribution along the ramp and the cowl surface. On following the ramp surface (red line) a plateau region is observed which depicts the constant pressure along the ramp surface. The sudden rise in pressure at X 90mm is due to the cowl tip shock. The peaks are observed in the regions of shock-shock interactions. Alternate crests and troughs evident of the oblique shock train in the isolator. The pressure distribution on the cowl surface (black line) follows the same profile of the ramp surface. The pressure contours for cowl angles 4, 5 and 6 deg are found to be similar with structure of recompression shock train. In these cases, the cowl shock and the ramp shock interact significantly thereby affecting the pressure rise at throat region. The tabulation of pressure values at the throat region with other performance parameters will be mentioned shortly in the report. The velocity contours at the isolator exit are an array of various individual velocity values. This non-uniformity is caused by the phenomenon call as Flow Distortion. Flow distortion in a scramjet inlet creates unsteadiness and hampers its pressure recovery. The variation of flow distortion with respect to the cowl angle is discussed further in this section. Contours of velocity profile are plotted below to study the effect of ramp and cowl shock interaction on the flow field of the isolator. Table 2 Variation of performance parameters with cowl angle From Table 2 it can be seen that as the cowl angle increases, the pressure recovery decreases due to increase in the flow distortion. The flow distortion increases due to increase in the strength of the incident shock due to interaction of the former with the expansion waves emanating from the ramp shoulder. The inlet is in the start condition for all the cowl angles as Mach number at the isolator end is greater than 1. The plots of the performance parameter are plotted below editor@iaeme.com

10 Senthilkumar P, Sharan Chandran M, Sreeja Sadasivan Govindha rasu N and K Senthil kumar Figure 12 Static Pressure vs Cowl angle Figure 13 Mach number vs Cowl angle Figure 14 Pressure Recovery vs Cowl Angle Figure 15 Flow distortion vs Cowl Angle As depicted in Figs 12 to 15 above, increase in cowl angle decreases the strength of the incident shock. Separation regions are lesser in the final three cases as the boundary layer is resistant to the adverse pressure (incipient separation shock). But flow distortion increases with cowl angle as well which leads to inlet unsteadiness and loss in pressure recovery as shown above. 5. CONCLUSION The effect of cowl angle is studied on the performance of the scramjet engine. A scramjet inlet prototype is designed in Solid Works followed by CFD analysis in ANSYS Fluent 15.0 to study the effects of cowl angle static pressure rise, total pressure recovery and Flow Distortion. It is noticed that the increase in cowl angle increases the flow distortion. This leads to inlet unsteadiness and loss in total pressure recovery. On the other hand, the incident shock strength decreases so that the separation region becomes lesser in size editor@iaeme.com

11 Effect of Cowl Angle in the Performance of Scramjet Air Intakes REFERENCES [1] Saied Emami, Carl A. Trexler, Aaron H. Auslender and John P. Weidner, Experimental Investigation of Inlet-Combustor Isolators for a Dual-Mode Scramjet at a Mach Number of 4, NASA Technical Paper [2] Scott D Holland, Internal Aerodynamics of a Generic Three Dimensional Scramjet Inlet at Mach10, NASA Technical Paper [3] Henckels, A., and P. Gruhn. Study on aerothermal effects of viscous shock interaction in hypersonic inlets. Fifth European Symposium on Aerothermodynamics for Space Vehicles. Vol [4] Ran, Hongjun, and Dimitri Mavris. Preliminary design of a 2D supersonic inlet to maximize total pressure recovery. AIAA Paper , Sept 2005 [5] Sivakumar, R., and V. Babu. Numerical simulations of flow in a 3-D supersonic intake at high Mach numbers. Defence Science Journal 56.4 (2006): [6] Krause, M., B. Reinartz, and J. Ballmann. Numerical computations for designing a scramjet intake. Department of mechanics, RWTH Aachen University. 25th International congress of aeronautical sciences [7] Janarthanam, S. and Babu, V. (2012). Numerical simulations of the flow through the inlet and isolator of a mach 4 dual mode scramjet. The Aeronautical Journal 116(1182), [8] Saha, Soumyajit, and Debasis Chakraborty. Hypersonic Intake Starting Characteristics A CFD Validation Study. Defence Science Journal 62.3 (2012): [9] Manoviraj Singh S, Arjun Singh, Govindha Rasu N, Performance Improvement Study of CESSNA-172 Aircraft using CFD, International Journal of Mechanical Engineering and Technology 8(9), 2017, pp [10] Senthilkumar P, Sreeja Sadasivan, A review on high speed flow with micro ramps, Journal of Chemical and Pharmaceutical Sciences, Special issue-3, 2016, [11] Senthilkumar P, Kamalraj Thiyagaraja, Numerical simulation of a scramjet intake with micro-cavity, Journal of Chemical and Pharmaceutical Sciences, Special issue-3, 2016, [12] Rajesh Prabha N, Edwin Raja Dhas J and Ramanan G, Finite Element Structural Analysis of Connecting Rod of Aa7075-Tic Composite Using ANSYS, International Journal of Mechanical Engineering and Technology 8(7), 2017, pp [13] Yousif Khudhair Abbas, Effect of Sideslip Angle on the Balance of Aircraft Moments Through Steady - State Spin. International Journal of Civil Engineering and Technology, 8(10), 2017, pp [14] Ali Salah Ameen, Dr. Ajeet Kumar Rai, Analysis of Electronic Chips Microchannel by Using Ansys Software, International Journal of Advanced Research in Engineering and Technology (IJARET) ISSN (Print) ISSN (Online) Volume 5, Issue 7, July (2014), pp editor@iaeme.com