scholarly journals Effect of Micro-Vortex Generator in Hypersonic Inlet

2011 ◽  
pp. 10-13
Author(s):  
Vivek V. Kumar ◽  
Surendra Bogadi
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Vortex Generator ◽  
Adverse Pressure Gradient ◽  
Actual Case ◽  
Hypersonic Inlet ◽  
Hypersonic Speeds ◽  
Structured Meshes ◽  
Free Stream Mach Number ◽  
Micro Vortex Generator

In the present study computational tests were carried out to get an understanding of the flow field in a pure mixedcompression hypersonic inlet at a free stream Mach number of 7 and an altitude of 35km. Structured meshes have been used for depicting the motion of fluid inside the inlet. First, a grid has been selected after conducting a grid study. Two dimensional simulations were carried out with standard sst k-ω model using FLUENT. Computational results are compared with the available data. The results obtained from the computational tests revealed several important flow field details at hypersonic speeds. The basic shock structure inside the inlet was obtained. The boundary layer formed inner side of the engine had an adverse pressure gradient on the top ramp. Due to this the boundary layer thickens and the static pressure starts to decrease whose effect leads till the trailing edge of inlet. By providing small wedge shaped Micro-Vortex Generator (MVG) where the shockboundary layer occurs we can smooth the boundary layer formed inside the inlet. Thus there will be more efficient compression than the actual case. The results obtained in the present series of tests, could help the hypersonic inlet design optimization at offdesign condition

scholarly journals Experimental Investigation of the Vortex Flow around an Axisymmetric Body with the Five-Hole Probe and Hot Wire Probe

2020 ◽  
Author(s):  
amir farajollahi ◽  
Mojtaba Dehghan Manshadi ◽  
Kazem Hejranfar
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Incidence Angle ◽  
Vortex Generator ◽  
Adverse Pressure Gradient ◽  
Axisymmetric Body ◽  
The Body ◽  
Vortical Flow ◽  
Hot Wire ◽  
Wire Probe

Abstract Anaxisymmetric body experiences the vertical flow around itself at incidence angle. If the adverse pressure gradient is significant, the boundary layers separated and a vortex is formed. The flow over a submarine at AOA (angle of attack) has specified separation of boundary layer and large vortex structures around the body. This flow influences body drag, acoustic and maneuverability. A propermethod to decrease and control the impacts of this separated flowis to use vortex generator. The mainobjective of the present study is to investigate the flow field around a Suboff model with applying the vortex generator by using the hot wire and five-hole probe in 0° ≤ α ≤ 20° angles of attack. The novelty of present study is application of two experimental method, (hot wire probe and five-hole probe) which can help to precisely study the structure of three-dimensional vortical flow field, the boundary layer velocity profiles and probability of the separation on the model with and without existence of vortex generator. The results indicate that vortex generators significantly decrease cross-flow separation, the size of vortices and the vortical flow.

Micro vortex generator control of axisymmetric high-speed laminar boundary layer separation

Shock Waves ◽  
2014 ◽  
Vol 25 (5) ◽  
pp. 521-533 ◽  
Author(s):  
D. Estruch-Samper ◽  
L. Vanstone ◽  
R. Hillier ◽  
B. Ganapathisubramani
Keyword(s):  
Boundary Layer ◽  
Laminar Boundary Layer ◽  
High Speed ◽  
Vortex Generator ◽  
Boundary Layer Separation ◽  
Layer Separation ◽  
Micro Vortex Generator

scholarly journals Potential of Micro-Vortex Generators in Enhancing the Quality of Flow in a Hypersonic Inlet-Isolator

2020 ◽  
Vol 77 (1) ◽  
pp. 1-10
Author(s):  
Azam Che Idris ◽  
Mohd Rashdan Saad ◽  
Konstantinos Kontis
Keyword(s):  
Boundary Layer ◽  
Flow Separation ◽  
Structural Damage ◽  
Vortex Generator ◽  
Transition To Turbulence ◽  
Cruise Missile ◽  
Passive Flow Control ◽  
Attachment Point ◽  
Hypersonic Inlet ◽  
The Impact

The rush to be the first to demonstrate a practical hypersonic cruise missile has never been more frantic among the world’s superpowers, especially since China and India have also announced their own programme. The main hurdle for safe hypersonic flight is the severe shock wave-boundary layer interaction (SWBLI) that could induce flow separation. The separation could lead to inlet unstart and also structural damage at the flow re-attachment point. The simplest method to control these phenomena is by using passive flow control devices such as micro-vortex generator (MVG). The MVG is typically sized in the range of sub-boundary layer and the vortex generated can induce an early transition to turbulence thus avoiding or reducing the impact of flow separation. Many studies have been published with regard to MVG, but most were done in low supersonic speed and not in the hypersonic flow regime. In the current study, the MVG array was placed strategically at various locations on a hypersonic inlet-isolator representative geometry. The MVG has been proven to be very effective in eliminating or reducing the size of flow separation thus reducing the associated peak pressure at the re-attachment point.

Impact of a Combination of Micro-Vortex Generator and Boundary Layer Suction on Performance in a High-Load Compressor Cascade

2018 ◽  
Author(s):  
Shan Ma ◽  
Wuli Chu ◽  
Haoguang Zhang ◽  
Chuanle Liu
Keyword(s):  
Boundary Layer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Vortex Generator ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Control Methods ◽  
Compressor Cascade ◽  
Boundary Layer Suction ◽  
Micro Vortex Generator

The performance of a compressor cascade is considerably influenced by flow control methods. In this paper, the synergistic effects of combination between micro-vortex generators (MVG) and boundary layer suction (BLS) are discussed in a high-load compressor cascade. Seven cases, which are grouped by a kind of micro-vortex generator and boundary layer suction with three locations, are investigated to control secondary flow effects and enhance the aerodynamic performance of the compressor cascade. The MVG is mounted on the end-wall in front of the passage. The rectangle suction slot with three radial positions is installed on the blade suction surface near the trailing edge. The numerical results show that: at the design condition, the total pressure loss is effectively decreased as well as the static pressure coefficient increase when the combined MVG and SBL method (COM) is used, which is superior to MVG in an aerodynamic performance. At the stall condition, the induced vortex coming from MVG could mix the low-energy fluid and mainstream, which result in the reduced separation, and the total pressure loss decreased by 11.54% when the suction flow ratio is 1.5%. The total pressure loss decreases by 14.59% when the COM control methods are applied.

scholarly journals Numerical Study on the Effect of an Off-Surface Micro-Rod Vortex Generator Placed Upstream NACA0012 Aerofoil

2021 ◽  
Vol 321 ◽  
pp. 01011
Author(s):  
Abderrahim Larabi ◽  
Michael Pereira ◽  
Florent Ravelet ◽  
Tarik Azzam ◽  
Hamid Oualli ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Passive Control ◽  
Numerical Study ◽  
Vortex Generator ◽  
Adverse Pressure Gradient ◽  
Control Device ◽  
Local Flow ◽  
Moderate Reynolds Number ◽  
Flow Deceleration ◽  
Drag Ratio

In this paper, 3D numerical simulations have been carried out to enhance the understanding of a flow over a passive control device composed of micro cylinder with, d/c = 1.34% placed in the vicinity of NACA0012 aerofoil wing, by means of γ–Reθt transition sensitive turbulence model meant to predict the separation induced by transition achieved for aerofoils operating at moderate Reynolds number (Re = 4.45×105). Results show that the separation of the boundary layer has been eliminated by the passive static vortex generator at stall regime due to the injection of free-stream momentum to the boundary layer. The early transition to turbulent state overcomes the local flow deceleration of an adverse pressure gradient and remains sticked to the wall the boundary layer. Furthermore, the wing aerodynamic performance are improved as drag is reduced and lift is enhanced which is straight forward linked to the lift to drag ratio gain that varies from 22.68% to 134.17% at post stall angles of attack.

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