Design and Evaluation of Generic Bump for Flow Control in a Supersonic Inlet Isolator

2021 ◽  
Vol 143 (5) ◽  
Author(s):  
Md. Raihan Ali Khan ◽  
A. B. M. Toufique Hasan
Keyword(s):  
Mach Number ◽  
Flow Control ◽  
Performance Improvement ◽  
Separation Bubble ◽  
Geometric Constraints ◽  
General Shape ◽  
Supersonic Inlet ◽  
Internal Shock ◽  
Boundary Layer Interaction ◽  
Wave Boundary

Abstract In this paper, the geometry of a supersonic inlet isolator is modified by the introduction of a two-dimensional (2D) bump to control the complex lip shock wave/boundary layer interaction (SWBLI). The bump is of general shape whose profile is designed primarily based on the inviscid theory of oblique shock waves, which accommodates the effect of freestream conditions; particularly, the flow Mach number. Further, the geometric constraints of the inlet are taken into consideration to generate a contoured bump. This well-designed generic bump is tested in the range of flight Mach number of 2.5 to 3.8 through numerical computations. The adopted computational methods are validated with the available experimental data. Results showed that the modified inlet using the present generic bump changes the internal shock structure, weakens the intensity of SWBLI, and subsequently reduces shock reflection phenomena which are prevalent in baseline inlet. The wall characteristics such as separation bubble (SB), skin friction, and total pressure loss are found to be reduced in inlet with bump. The SB in baseline inlet typically corresponds with the geometric profile of the bump. As a result, ramp of baseline inlet is apparently replaced by this generic bump, which eliminates the low momentum fluid adjacent to the wall and the passage of modified inlet is found to be mostly occupied by high momentum supersonic flow. The flow control and associated performance improvement are linked with this modification of supersonic inlet isolator.

Progress in Hypersonic Inlet Flow Controls by Magnetohydrodynamic

2015 ◽  
Vol 730 ◽  
pp. 311-315
Author(s):  
Shuai Zhang ◽  
Guang Hong ◽  
Dong Yu Zhang ◽  
Fu Qianag Cheng
Keyword(s):  
Flow Control ◽  
Large Scale ◽  
Leading Edge ◽  
Mhd Flow ◽  
Surface Flow ◽  
Near Surface ◽  
Hypersonic Inlet ◽  
Boundary Layer Interaction ◽  
Large Scale Flow ◽  
Wave Boundary

The magnetohydrodynamic (MHD) flow control techniques for inlet based on “AJAX” vehicle concept have attracted many researchers’ attention for its potential wide applications in the supersonic flights or the hypersonic ones. In this paper, a preliminary discussion for basic problems that are ionization and magnetic field in the MHD applications is presented. And three major MHD techniques for enhancing inlet aerodynamic performance and operability, which are large scale flow control for enlarging flight scope in Mach/angle-of-attack, near surface flow control for shock wave/boundary layer interaction and leading edge heating-transfer control for enormous thermal load on leading edge of the compression ramp, are briefly explained and reviewed.

scholarly journals Effect of wall temperature on separation bubble size in laminar hypersonic shock/boundary layer interaction flows

2019 ◽  
Vol 11 (11) ◽  
pp. 168781401988555 ◽  
Author(s):  
Amjad A Pasha ◽  
Khalid A Juhany
Keyword(s):  
Boundary Layer ◽  
Heat Flux ◽  
Mach Number ◽  
Wall Temperature ◽  
Separation Bubble ◽  
Layer Interaction ◽  
Shock Boundary Layer Interaction ◽  
Boundary Layer Interaction ◽  
Hypersonic Shock ◽  
Bubble Length

At hypersonic speeds, the external wall temperatures of an aerospace vehicle vary significantly. As a result, there is a considerable heat transfer variation between the boundary layer and the wall of the hypersonic vehicle. In this article, numerical computations are performed to investigate the effect of wall temperature on the separation bubble length in laminar hypersonic shock-wave/boundary-layer interaction flows over double-cone configuration at the Mach number of 12.2. The flow field is described in detail in terms of different shocks, expansion fans, shear layer and separation bubble. The variation of the Prandtl number has a negligible effect on the flow field and wall data. A specific heat ratio of less than 1.4 results in the better prediction of wall pressure and heat flux in the shock/boundary-layer interaction region. It is observed that as the wall temperature is increased, the separation bubble size and hence the separation shock length increases. The high firmness of the laminar boundary-layer at a high Mach number shows that the wall temperature in the shock/boundary-layer interaction region has little effect. The peak wall pressure and heat flux decrease with an increase in wall temperature. An estimation is developed between separation bubble length and wall temperature based on the computed results.

Perforated Wall in Controlling the Separation Bubble Due to Shock Wave –Boundary Layer Interaction

2019 ◽  
Vol 0 (0) ◽  
Author(s):  
K. Raja sekar ◽  
S. Jegadheeswaran ◽  
R. Kannan ◽  
P. Manigandan
Keyword(s):  
Numerical Simulation ◽  
Boundary Layer ◽  
Shock Wave ◽  
Separation Bubble ◽  
Wedge Angle ◽  
Layer Interaction ◽  
Wave Boundary Layer ◽  
Perforated Wall ◽  
Boundary Layer Interaction ◽  
Wave Boundary

Abstract The shock wave boundary layer interaction (SWBLI) induced separation bubble formation (SB) and its control has been investigated numerically in the mixed compression type of intake in the scramjet engine. The external compression has occurred due to the three successive oblique shocks formed from the three successive ramps of the forebody with the semi-wedge angle of 7.6, 7.0, and 9.4 respectively. The intake is designed in such a way that all three shocks converge and impinge on the leading edge of the cowl lip for the operating Mach number of 5.0. The numerical simulation is carried out by solving steady, compressible 2-D RANS equations using transitional SST k-ω turbulence model to capture the influence of SWBLI in the performance of supersonic intake. The formation of SB and its control by establishing the perforated wall in its proximity are investigated for three different cases based on the perforation with respect to SSB. Findings of the numerical simulation have concluded that the size of the SB decreases to an acceptable level while establishing the perforation in the entire fore-body wall in the isolator region. The feedback loop established between the upstream and downstream of SB could be a possible reason for reducing its size.

Flow Diagnostics in Shock Wave-Boundary Layer Interaction Experiments in Hypersonic Flow

2014 ◽  
Vol 660 ◽  
pp. 669-673
Author(s):  
Mohd Rashdan Saad ◽  
Azam Che Idris ◽  
Konstantinos Kontis
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Flow Control ◽  
High Speed ◽  
Vortex Generators ◽  
Flow Diagnostics ◽  
Control Techniques ◽  
Layer Interaction ◽  
Boundary Layer Interaction ◽  
Wave Boundary

Shock Wave-Boundary Layer Interaction (SBLI) is a phenomenon occurring in high-speed propulsion systems that is highly undesirable. Numerous methods have been tested to manipulate and control SBLI which includes both active and passive flow control techniques. To determine the improvements brought by the flow control techniques, advanced and state-of the-art flow diagnostics and experimental techniques are required, especially when it involves high-speed flows. In this study, a number of advanced flow diagnostics were employed to investigate the effect of micro-vortex generators in controlling SBLI in Mach 5 such as Pressure Sensitive Paints (PSP), Particle Image Velocimetry (PIV), schlieren photography and oil-flow visualization. The flow diagnostics successfully visualized the boundary layer separation and also the improvements brought by the micro-vortex generators.

scholarly journals Effect of Varying Ramp Angle and Leading-Edge Bluntness on the Behavior of Ramp Induced Shock Wave over Triple Ramped Cone Flare Configuration at Hypersonic Speed

2019 ◽  
Vol 8 (3) ◽  
pp. 8127-8138
Keyword(s):  
Shock Wave ◽  
Bubble Size ◽  
Separation Bubble ◽  
Flow Characteristics ◽  
Engine Performance ◽  
Leading Edge ◽  
Experimental Conditions ◽  
Boundary Layer Interaction ◽  
Induced Flow ◽  
Wave Boundary

Numerical simulation results are presented to show the effect of ramp angle variations and leading-edge bluntness on the flow around triple ramped cone flare in hypersonic flow. This study investigates the changes associated with shock wave boundary layer interaction due to ramp induced flow breakdown and the fluctuation in flow in the presence of blunted leading edge. This type of ramp junctions typically features in re-entry vehicles, engine intakes, system and subsystem junctions, control surfaces, etc. Ramp junctions usually are associated with strong separation bubble that has significant upstream influence impacting the effectiveness of aerodynamic surfaces, engine performance, thermal behavior and stability. Computation studies are carried out using finite volume-based RANS solver, accuracy of second order and considering compressible laminar flow characteristics, with solver settings provided similar to experimental conditions as per literature. Comprehensive double ramp studies with suggestions on reducing the separation bubble size are invariantly considered in literature, however there has been no study in understanding the inclusion of additional ramps in such flow scenarios, hence efforts are taken to understand the benefits and implications of including a third ramp along with varying bluntness on the bubble size and its upstream intensity

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