Pressure Distribution Measurements in Scramjet Isolators under Asymmetric Supersonic Flow

2006 ◽  
Author(s):  
Cheng-peng Wang ◽  
Kun-Yuan Zhang ◽  
Ke-Ming Cheng
Keyword(s):  
Supersonic Flow ◽  
Pressure Distribution

scholarly journals Aerodynamic Layout Designs Using of Microflaps for Flow Control of a Supersonic Finned Projectile

2019 ◽  
Vol 37 (2) ◽  
pp. 344-353
Author(s):  
Xiaopeng Li ◽  
Shun Kang ◽  
Tianming Wang ◽  
Yangzheng Zhao
Keyword(s):  
High Pressure ◽  
Control System ◽  
Supersonic Flow ◽  
Flow Control ◽  
Pressure Distribution ◽  
Surface Pressure ◽  
Pressure Zone ◽  
Bow Shocks ◽  
Flow Control System ◽  
Interference Mechanism

In this paper, the flow control system consists of some small microflaps located between the rear fins of the projectile. These microflaps can alter the flow field in the finned region of the projectile resulting in asymmetric pressure distribution and thus producing control forces and moments, furthermore to provide directional control for a supersonic projectile. Due to the small size and high speed characteristics of projectile, which is with fast and valid response characteristics, this flow control system has initially shown excellent potential in terms of supersonic flow control. The CFD simulation used here solves steady-state Reynolds-averaged Navier-Stokes equation with two-equation turbulence model k-ε. Firstly, we investigate the flow mechanism around microflap in supersonic flow, the flow fields around the microflap are complex, involving three-dimensional shock-shock, shock-boundary layer interactions. Secondly, for the microflap and the fin of Basic Finner configuration, the influence of microflap geometric parameters, microflap locations on aerodynamics is obtained and the interference mechanism is explored. Finally, several typical roll and pitch control layouts are described. According to the simulation results and their analysis, some preliminary conclusions can be drawn: by analyzing the flow interference mechanism between microflap and the fin, we find that the separated shocks ahead of the microflap, the bow shocks around microflap, and the trailing-edge wake, have influences on fin's surface pressure; among these factors, the bow shocks are stronger than separated shocks, furthermore it can generate larger high pressure region. Then we find out the aerodynamic characteristics of several typical control layouts at a supersonic speed, Ma=2.5, furthermore, hence nearly 4.8% drag is increase compared with the condition without microflap. As the number of microflaps increasing, the control aerodynamic forces and moments increases almost linearly. With a proper layout of the microflap's location, quick change in the surface pressure distribution can be achieved for rear fins of the projectile, the microflap should be mounted that can increase the high pressure zone, meanwhile, reduce the low pressure zone on the surface of fins, thus modulating the projectile's attitude can be realized.

Calculation of the Flow Properties Up- and Downstream of and Within a Supersonic Turbine Cascade

1972 ◽  
Author(s):  
O. K. Lawaczeck
Keyword(s):  
Supersonic Flow ◽  
Wind Tunnel ◽  
Pressure Distribution ◽  
Theoretical Method ◽  
Experimental Results ◽  
Turbine Cascade ◽  
Flow Properties ◽  
Flow Through

This paper deals with the flow through cascades, which have a supersonic flow up- and downstream. A theoretical method is described to calculate, under certain conditions, the flow quantities up- and downstream of and within the cascade and to determine the pressure distribution on the blades. The method is compared with experimental results, carried out in the cascade wind tunnel of the Aero dynamische Versuchsanstalt Göttingen (AVA), Germany.

Supersonic flow past a bluff body with a detached shock Part II Axisymmetrical body

1956 ◽  
Vol 1 (5) ◽  
pp. 490-496 ◽  
Author(s):  
W. Chester
Keyword(s):  
Mach Number ◽  
Supersonic Flow ◽  
Pressure Distribution ◽  
Bluff Body ◽  
The Body ◽  
Adiabatic Index ◽  
Radius Of Curvature ◽  
Supersonic Stream ◽  
High Mach Number ◽  
High Mach

This paper considers the flow at high Mach number behind the curved shock formed when a supersonic stream impinges on an axisymmetrical body with a rounded nose.The solution is obtained as a double expansion in $\delta = (\gamma -1)|(\gamma +1),$ where γ is the adiabatic index, and M−2; the expansion is developed to within terms of order (δ+M−2)3.Expressions are obtained for the distance between the body and the shock, the radius of curvature of the shock compared with that of the body, and the pressure distribution on the body.

A Note on Supersonic Wing Integral Equations in Unsteady Flow

1952 ◽  
Vol 3 (4) ◽  
pp. 294-296 ◽  
Author(s):  
John W. Miles
Keyword(s):  
Supersonic Flow ◽  
Unsteady Flow ◽  
Velocity Distribution ◽  
Pressure Distribution ◽  
Integral Equations ◽  
Thin Wing

SummaryThe integral equations for the pressure distribution on an oscillating thin wing having a prescribed velocity distribution in a supersonic flow are developed.

Supersonic Flow over Thin Symmetrical Wings with Given Surface Pressure Distribution

1952 ◽  
Vol 3 (4) ◽  
pp. 263-279 ◽  
Author(s):  
F. A. Goldsworthy
Keyword(s):  
Boundary Condition ◽  
Inverse Problem ◽  
Integral Equation ◽  
Supersonic Flow ◽  
Pressure Distribution ◽  
Surface Pressure ◽  
Integral Equation Method ◽  
Normal Velocity ◽  
Vertical Derivative ◽  
Surface Pressure Distribution

SummaryThe inverse problem of determining the supersonic flow past a thin symmetrical wing at zero incidence with given surface pressure distribution is solved for wings of arbitrary plan form. Expressions are obtained for the vertical derivative of the potential on the wing surface from which, using the linearised boundary condition of zero normal velocity at the surface, the profile of the wing can be designed. The integral equation method adopted by J. C. Evvard and extended by G. N. Ward is used. The analysis cannot be applied to pointed wings, whose leading edges are subsonic. The results in Part I are completely general and are applied to specific problems in Part II.

Sign in / Sign up

Export Citation Format

Share Document