Shock Wave Turbulent Boundary Layer Interaction by ISERP ISERP

ASMELASH HAFTU AMAHA* et al. / (IJAEST) INTERNATIONAL JOURNAL OF ADVANCED ENGINEERING SCIENCES AND TECHNOLOGIES Vol No. 5, Issue No. 2, 156 - 161

Shock Wave Turbulent Boundary Layer Interaction Department of Mechanical Engineering Defence Institute of Advanced Technology, Girinagar, Pune-411025, Maharashtra, India E-mail: ahaftu@yahoo.com

I. INTRODUCTION The interaction between a shock wave and a turbulent boundary layer or (SWTBLI) remains one of the most outstanding problems of modern high speed fluid dynamics. The complicated nature of the interaction embodies most challenging effects and raises difficult problems which are still largely unsolved both in internal and external flows. SWTBLIs are prevalent in a variety of high speed applications, such as long haul civil transportation systems, as well as other high-speed manoeuvring vehicles, and thus can induce separation which causes loss of control surface effectiveness, loss of total pressure, drop of air intake efficiency and may be at the origin of large scale fluctuations such as detrimental air intake buzz or engine extinction, buffeting or fluctuating side loads in separated propulsive nozzles. Therefore, an in-depth knowledge of the phenomena is essential for efficient aerodynamic and propulsion design. [1-3, 5-8]

Abstract: - A shock wave turbulent boundary layer interaction has been analyzed computationally in a twodimensional compression ramps for a free stream Mach number of 2.85. Ramp angles ranging from 8 to 240 were used to produce the full range of possible flow fields, including flows with no separation, moderate separation, and significant amount of separation. The model has been studied based on a commercially available Computational Fluid Dynamics (CFD) Code that employs Shear Stress Transport (SST) of k-Ď&#x2030; turbulence model and Realizable k-É&#x203A; model. The CFD code and the turbulence models used are validated by comparing with experimental results available in literature. The computed data for surface pressure distribution and skin friction coefficient indicated a good comparison with the experiment. Numerical results obtained through the present series of computations indicate an increased separation and reattachment locations, when compared to experiment. Measurements of total pressure have shown to increase from separation to reattachment points and increase when ramp angle increases.

AMARJIT SINGH

Department of Aerospace Engineering, Defence Institute of Advanced Technology, Girinagar, Pune-411025, Maharashtra, India E-mail: amarjit100@gmail.com

ASMELASH HAFTU AMAHA

Keywords:- Ramp flow, Shock wave, Separation, Reattachment, turbulent boundary layer.

A review of literature published in this area indicates that, the complex interaction between shock waves and turbulent boundary layers continue to be of great interest to researchers. There is a need for additional research work. Some of the authors who have extensively studied the flow field in the ramp flow include: (Delery (1985) [1]), (Daniel Arnal and J.M. Delery (2004) [2]), (A.B. Oliver et al (2007) [3]), (Settles et al (1994) [4]), and (D. W. Kuntz [5]).

ISSN: 2230-7818

Page 156

ASMELASH HAFTU AMAHA* et al. / (IJAEST) INTERNATIONAL JOURNAL OF ADVANCED ENGINEERING SCIENCES AND TECHNOLOGIES Vol No. 5, Issue No. 2, 156 - 161

II.

1. The ramp upstream influence, hence upstream influence length Lo, has considerably increased. 2. A first shock associated with separation forms well upstream of the ramp. 3. A second shock originates from the reattachment region on the ramp which intersects the separation shock at a short distance from the wall.

The investigation described in this report was conducted in order to provide detailed physical description of the SWTBLIs to improve the capability to predict the surface pressure and wall shear stress for these interactions. The investigation was conducted using the CFD code, and thus has produced new information concerning loss of total pressure ratio in the separated zone, which adversely affects the performance of high speed aeronautical/aerospace devices.

pressure rise is high enough to induce separation of the boundary layer (Figure 1). In this situation:

THEORY

Compression Ramp Flow

Shock wave turbulent boundary layer interactions are basically two-dimensional and three-dimensional. 2-D interactions include: the ramp flow, the impinging reflecting shock, and the pressure discontinuity resulting from adaptation to a higher downstream pressure level. 3-D interactions include, swept shock boundary layer interaction. In this paper, Ramp flow interaction, which corresponds to a control surface or an air-intake compression ramp, has been considered.

The ramp flow interaction occurs when supersonic flow along a flat plate is compressed by a wedge or ramp of angle α. When the ramp angle α is small, the overall flow structure is not much affected by the interaction taking place at the ramp origin. The main difference is a spreading of the wall pressure distribution, the step of the inviscid solution being replaced by a progressive rise between the upstream pressure level p0 and the final value p1 corresponding to the oblique shock equations. When the ramp angle α is increased (hence the shock strength), the upstream influence distance (L0 defined as the distance between the interaction onset and the ramp origin) increases accordingly and a situation can be reached where the

Fig1. The structure of a ramp flow with boundary layer separation [2]

III.

NUMERICAL SIMULATION

Grid generation and boundary conditions Commercially available CFD software has been used to simulate twodimensional, high-speed, turbulent flow with air as the fluid for a flow over a compression-ramp at free-stream Mach numbers (M0= 2.85). The flow conditions and simulation setups used are given in this section. To study the flow-field of the interaction, the basic ramp flow geometry used in [3] has been adopted for the present study, primarily due to the availability of experimental data. Basic mesh details of the ramp (200) are shown in figure 2 with a typical grid distribution adopted near the corner region.

ISSN: 2230-7818

Page 157

ASMELASH HAFTU AMAHA* et al. / (IJAEST) INTERNATIONAL JOURNAL OF ADVANCED ENGINEERING SCIENCES AND TECHNOLOGIES Vol No. 5, Issue No. 2, 156 - 161

M∞

T0[K]

P0[Pa]

δ [mm]

Re∞[m-1]

2.87

280

6.9E+05

60.8.106

120

2.87

280

6.9E+05

60.8.106

160

2.85

268

6.9E+05

65.6.106

200

2.85

258

6.9E+05

69.5.106

240

2.84

262

6.9E+05

68.3.106

Ramp

Table 2. Reference values used in data reduction Ramp

160

200

240

δRef[mm]

23561

The software is commercially available software that solves steady state and time accurate CFD problems on structured/unstructured grids. The commercial software has several solvers for both incompressible and compressible flows in both an implicit and explicit numerical framework.

Table1. Free stream and inflow boundary layer data

Computations were made by using grid to grid file interpolation technique with three different grids [Grid 1 (63000 cells), Grid 2 (87000 cells) and Grid 3 (101500 cells)] in two blocks and 0 < y+ < 0.4. Density based implicit solver with upwind second order discretisation scheme for flow and transport equations was chosen to capture shocks better at cell faces. Convergence criteria were considered when residuals have fallen by three orders and measurement of shear stress at different locations converged to four significant digits all achieved after 30000 iterations.

PRef[Pa]

IV.

RESULTS AND DISCUSSIONS

Results of the simulation computed by the CFD Code are presented in this section along with discussion. Nondimensionalizing reference values are given in table2. Density contour plots and plots of skin friction coefficient, surface pressure distribution, and total pressure are presented.

Fig 2. Grid system (computational domain)

22856

The boundary conditions for the 2D ramp flow grid systems are pressure inlet at the inflow boundary, pressure_far_field on the boundary opposite the viscous wall (isothermal), and pressure outlet at the outflow. A convergence criterion for the residuals of all equations has been set to (10-6). Turbulent Intensity of (5%) and viscosity ratio of 5 has been used. Boundary layer thickness (δ0) computed from simulation was 25mm.

Fig 3. Plot of density contours near the corner region (80-240)

ISSN: 2230-7818

Page 158

ASMELASH HAFTU AMAHA* et al. / (IJAEST) INTERNATIONAL JOURNAL OF ADVANCED ENGINEERING SCIENCES AND TECHNOLOGIES Vol No. 5, Issue No. 2, 156 - 161

shows the region of rversed flows and the idea is as described by the density contour plots.

Fig. 6. Skin friction coefficient (200)

In figure 6, skin friction coefficient (Cf) computed by Realizable k-É&#x203A; and SST is compared to experiment. Both models show larger separation region compared to the experimental data. SST shows separation earlier and reattaches later. R. k-É&#x203A; is closer to the experiment in the recovery region

At 80, the flow is attached. Shock originates at the corner location and the curvature of the shock results from Mach number gradient in the incoming boundary layer. A greater upstream influence is seen in 160 ramp, showing the first evidence of separation line. The 200 and 240 corner angles indicate a clear significant separation. For 200 and 240, the intersection of separation and reattachment shocks indicates a slight bend from which a stronger shock appears due to merging of the two left running shocks. For this type of shock-shock interaction, the reflected wave and slip line are indicated in figure 1.

Fig. 4. Surface shear stress distribution on (80-160)

The shear stress distribution parallel to wall and corner surfaces for 80 to 160 is depicted in figures 4. In this case, only the 160 flow show the insipient separation condition.

Fig.7 Comparison of Surface pressure distributions on various compression ramp interactions (80-240) ramps

Fig. 5. Surface shear stress distribution on (16 -24 )

In figure 5, how the separated region increases strength is plotted for 160 angles. The negative wall

size of the with shock to 240 ramp shear stress

Figure 7 shows the static pressure ratios for various corner angles (shock strengths) in the flow field and on the ramp surfaces for flow at mach 2.85. The (80 and 120) show unseparated flows. Incipient separation is observed at 160 which show a shift in the sharp pressure rise indicating

ISSN: 2230-7818

Page 159

ASMELASH HAFTU AMAHA* et al. / (IJAEST) INTERNATIONAL JOURNAL OF ADVANCED ENGINEERING SCIENCES AND TECHNOLOGIES Vol No. 5, Issue No. 2, 156 - 161

agreement between themselves and to the experiment. However, when the interaction strength increases (200 and 240), the turbulence models show slightly higher separation size (figure 6 (b) and (c)) may be due to lower Reynolds number effects. As seen from figure 8, the SST model separates too early.

Fig.9. total pressure ratios at Sp, 0, and R-points for 240 ramp

the beginning of upstream influence. The pressure distributions for 200 and 240 indicate that in a region of separation, the wall pressure initially rises gradually through the unsteady motion of the separation shock. The wall pressure reaches a plateau in the fully separated region and then rises gradually again after the flow reattaches. The pressure distribution reveals the large upstream influence and large streamwise extent of the interaction.

In figures 9, Comparison of total pressure loss is presented. This comparison on three different locations of the interaction (at separation location „S‟ at interaction origin „0‟ at Reattachment location „R‟) show that the total pressure loss inside the boundary layer increases from separation to the reattachment points as the flow moves down stream of the interaction. This is because total pressure depends on static pressure which increases downstream and the effect of the separation is carried downstream.

Fig.8 Comparison of surface pressure distributions based on SST, R.k-ɛ and Experiment. (a) 160 ramp, (b) 200 ramp, (c) 240 ramp

For weaker interaction (figure 8 (a)) the turbulence models are in a closer

Fig.10. total pressure ratios 8-240 at interaction origin

ISSN: 2230-7818

Page 160

ASMELASH HAFTU AMAHA* et al. / (IJAEST) INTERNATIONAL JOURNAL OF ADVANCED ENGINEERING SCIENCES AND TECHNOLOGIES Vol No. 5, Issue No. 2, 156 - 161

CONCLUSIONS

Flow has been simulated for compression ramps of deflection angles ranging from 8-240 at M0 of 2.85. Using the density contours, incipient separation was observed at 160, and a fully separated situation has been observed explicitly showing the separation and reattachment shocks for higher deflection angles.

The skin friction coefficient and surface pressure distributions predicted by the CFD Code for the 2-D compression ramps based on SST and Realizable k-ɛ models have shown greater tendency to separate, possibly an effect of the low Reynolds number employed in the models. In both cases the SST model has predicted the separation region earlier. Total pressure loss has shown to increase from separation to reattachment points and increase when corner angle increases.

[2] Daniel Arnal and Jean Delery: Shock Wave Boundary Layer Interaction, NATO, May 2004. [3] A. B. Oliver et al: “Assessment of Turbulent Shock-Boundary Layer Interaction Computations Using the OVERFLOW Code,” School of Aeronautics and Astronautics Purdue University, January 2007. [4] Gary S. Settles and Lori j. Dodson: “Hypersonic Shock/Boundary Layer Interaction Database: New and Corrected Data,” NASA Contractor Report 177638, April 1994. [5] D. W. Kuntz et al: “Turbulent Boundary layer Properties Downstream of the Shock-Wave/ Boundary Layer Interaction,” AIAA journal Vol.25, No.5, May 1987. [6] K.Sinha, K. Mahesh et al: “Modelling the effects of shock unsteadiness in Shock turbulent Boundary Layer Interactions,” AIAA Journal Vol. 43, No. 3, March 2005. [7] Patrick Bookey et al.: “An experimental investigations of Mach 3 Shock-Wave Turbulent Boundary Layer Interactions,” Princeton, NJ08540, AIAA-2005-4899. [8] DC Wilcox: “Turbulence modelling for CFD,” 1994.M.

In figure 10, the loss in total pressure increases with increasing deflection angle, and thus the boundary layer requires more time to recover when the shock strength increases.

ACKNOWLEDGMENT

The authors are sincerely thankful to: (1). Prof. L.M. Patnaik, Vice Chancellor, DIAT (DU), (2). Prof. R.D. Misal, HOD, Department of Mechanical Engineering DIAT (DU), and (3). Computer Engineering Department of DIAT (DU), Girinagar, Pune-411025, Maharashtra, India, for supporting this work. REFERENCES [1] J.M. Delery: “Shock Wave Turbulent Boundary Layer Interaction and its Control,” Aerospace Sci. Vol. 22, pp. 209-280, 1985. 6

ISSN: 2230-7818

Page 161