The Static Stability of a Two-Dimensional Curved Panel in a Supersonic Flow, with an Application to Panel Flutter

1954 ◽  
Vol 21 (8) ◽  
pp. 556-565 ◽  
Author(s):  
Y. C. FUNG
Keyword(s):  
Supersonic Flow ◽  
Static Stability ◽  
Panel Flutter ◽  
Two Dimensional ◽  
Curved Panel

Evolution of weak shock wave in two-dimensional steady supersonic flow in dusty gas

2019 ◽  
Vol 160 ◽  
pp. 552-557 ◽  
Author(s):  
Rahul Kumar Chaturvedi ◽  
Pooja Gupta ◽  
L.P. Singh
Keyword(s):  
Shock Wave ◽  
Supersonic Flow ◽  
Weak Shock ◽  
Weak Shock Wave ◽  
Dusty Gas ◽  
Two Dimensional

Measurement of Recovery Factors and Friction Coefficients for Supersonic Flow of Air in a Tube: 2—Results Based on a Two-Dimensional Flow Model for Entrance Region

1952 ◽  
Vol 19 (2) ◽  
pp. 185-194
Author(s):  
J. Kaye ◽  
T. Y. Toong ◽  
R. H. Shoulberg
Keyword(s):  
Boundary Layer ◽  
Supersonic Flow ◽  
Laminar Boundary Layer ◽  
Flow Model ◽  
Variable Mass ◽  
Entrance Region ◽  
Two Dimensional ◽  
Dimensional Flow ◽  
Change Of State ◽  
Two Dimensional Flow

Abstract The first part of a program to obtain reliable data on the rate of heat transfer to air moving at supersonic speeds in a tube has been devoted to measurements made on adiabatic supersonic flow of air in a tube. The details of these measurements have been described in a previous paper. The calculated quantities such as the local apparent friction coefficient, recovery factor, Mach number, and so forth, were obtained from the simple one-dimensional flow model for which the properties of the stream are uniform at any section, and boundary-layer effects are ignored. The analysis of some of the same data given in the previous paper is undertaken here with the aid of a simplified two-dimensional flow model. The supersonic flow in the tube is divided into a supersonic core of variable mass with the fluid remaining in the core undergoing a reversible adiabatic change of state, and a laminar boundary layer of variable mass. The compressible laminar boundary layer increases in thickness in the direction of flow, and then undergoes a transition to a turbulent boundary layer. The two-dimensional flow model is limited here to the region where a laminar boundary layer appears to be present in the entrance region of the tube. The results of the analysis based on the two-dimensional flow model indicate that where the flow in the tube boundary layer appears to be laminar, the measured pressures and temperatures in the tube for adiabatic supersonic flow of air could have been predicted, with sufficient accuracy for engineering problems, from measured data for supersonic flow of air over a flat plate with a laminar boundary layer, and with zero pressure gradient.

Axially-symmetrical supersonic flow near the centre of an expansion

1950 ◽  
Vol 2 (2) ◽  
pp. 127-142 ◽  
Author(s):  
N.H. Johannesen ◽  
R.E. Meyer
Keyword(s):  
Approximate Solution ◽  
Supersonic Flow ◽  
Conventional Method ◽  
Velocity Gradient ◽  
Method Of Characteristics ◽  
Simple Wave ◽  
Two Dimensional ◽  
Perfect Gas ◽  
Initial Curvature ◽  
Boundary Wall

SummaryWhen a uniform, two-dimensional supersonic flow expands suddenly round a corner in a wall it forms a pattern known as a Prandtl-Meyer expansion or centred simple wave. If the flow is two-dimensional but not initially uniform, or if it is axially-symmetrical, the expansion is still centred, but is not a simple wave. An approximate solution is given in this paper for the isentropic, irrotational, steady two-dimensional or axially-symmetrical flow of a perfect gas in the neighbourhood of the centre of such an expansion. The solution is designed to replace the conventional method of characteristics in such a region.The main application is to a jet issuing from a nozzle that discharges into a container with a pressure lower than that in the nozzle; in particular, a formula is derived for the initial curvature, at the lip of the nozzle, of the boundary of the jet. The solution also applies to the flow near an edge in a boundary wall, and a formula is derived for the velocity gradient on the wall immediately downstream of the edge.

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