Numerical modeling of supersonic-flow deceleration in a two-dimensional duct in combustion of wall tangential hydrogen jets

1996 ◽  
Vol 32 (4) ◽  
pp. 399-405
Author(s):  
O. M. Kolesnikov
Keyword(s):  
Numerical Modeling ◽  
Supersonic Flow ◽  
Two Dimensional ◽  
Flow Deceleration

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

Numerical modeling of two-dimensional smoldering processes

1993 ◽  
Vol 95 (1-2) ◽  
pp. 170-182 ◽  
Author(s):  
M MOALLEMI ◽  
H ZHANG ◽  
S KUMAR
Keyword(s):  
Numerical Modeling ◽  
Two Dimensional

TRIGGERING OF DETONATION PROCESSES IN PROPULSION CHAMBER

2021 ◽  
Vol 14 (2) ◽  
pp. 40-45
Author(s):  
D. V. VORONIN ◽  
Keyword(s):  
Thermal Conductivity ◽  
Numerical Modeling ◽  
Gas Flow ◽  
Stokes Equations ◽  
Navier Stokes ◽  
Two Dimensional ◽  
Navier Stokes Equations ◽  
Chemically Reacting ◽  
Plane Disk ◽  
And Diffusion

The Navier-Stokes equations have been used for numerical modeling of chemically reacting gas flow in the propulsion chamber. The chamber represents an axially symmetrical plane disk. Fuel and oxidant were fed into the chamber separately at some angle to the inflow surface and not parallel one to another to ensure better mixing of species. The model is based on conservation laws of mass, momentum, and energy for nonsteady two-dimensional compressible gas flow in the case of axial symmetry. The processes of viscosity, thermal conductivity, turbulence, and diffusion of species have been taken into account. The possibility of detonation mode of combustion of the mixture in the chamber was numerically demonstrated. The detonation triggering depends on the values of angles between fuel and oxidizer jets. This type of the propulsion chamber is effective because of the absence of stagnation zones and good mixing of species before burning.

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.

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