Study of Supersonic Flowfield With Secondary Injection

2004 ◽  
Author(s):  
D. Sun ◽  
R. S. Amano
Keyword(s):  
Total Pressure ◽  
Boundary Layer Thickness ◽  
Stokes Equations ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Navier Stokes Equations ◽  
Weighted Essentially Nonoscillatory ◽  
Secondary Injection ◽  
Total Pressure Ratio

Two and three-dimensional steady flowfields generated by transverse secondary injection into a supersonic flow, was simulated by solving the Favre-averaged Navier-Stokes equations using the weighted essentially nonoscillatory (WENO) schemes. Both the two-dimensional and three-dimensional results are given. Some parameters affecting the penetration height and separation length of the interactive flowfield, including the total pressure ratio of the jet to the freestream, the boundary layer thickness, slot width, the Mach number of the freestream and injection, the jet angle, and the shape of the injection orifice in the 3D flowfield, were calculated in more detail.

Design and Experimental Study on the Surge Performance Improvement by the Effects of Bleed Slot Casing of a Centrifugal Compressor

2014 ◽  
Author(s):  
Hong Won Kim ◽  
Jae Hoon Chung ◽  
Hyo Seong Lee ◽  
Min Ouk Choi
Keyword(s):  
Centrifugal Compressor ◽  
Stokes Equations ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Design Parameters ◽  
Navier Stokes Equations ◽  
Operating Range ◽  
Surge Margin ◽  
Commercial Code

The primary design goal of a compressor is focused on improving efficiency. Secondary objective is to widen the compressor’s operating range. This paper presents a numerical and experimental investigation of the influence of the bleed slot to enlarge operating range for the 1.2MW class centrifugal compressor installed in a turbocharger. The main design parameters of the bleed slot casing are upstream slot position, inlet pipe slope, downstream slot position and width. The DOE (design of experiment) method was carried out to optimize the casing design. Numerical analyses were done by the commercial code ANSYS-CFX based on the three dimensional Reynolds-averaged Navier-Stokes equations. From the analysis, as the downstream slot position and width are smaller and upstream position is located away from impeller inlet, efficiency and pressure ratio are increased. Experimental works were done with and without the bleed slot casing. The simulation results were in good agreement with the test data. In case without the bleed slot casing, the surge margin value came out to be only 11.8% but with the optimized bleed slot design, the surge margin reached 23%. Therefore, the surge margin increase of 11.2% was achieved.

Investigation of Vortex Structures for Supersonic Jet Interaction Flowfield

2015 ◽  
Vol 798 ◽  
pp. 546-550 ◽  
Author(s):  
Asel Beketaeva ◽  
Amr H. Abdalla ◽  
Yekaterina Moisseyeva
Keyword(s):  
Stokes Equations ◽  
Pressure Ratio ◽  
Supersonic Jet ◽  
Three Dimensional ◽  
Horseshoe Vortex ◽  
Bow Shock ◽  
Navier Stokes ◽  
Navier Stokes Equations ◽  
Trailing Vortices ◽  
Barrel Shock

The three-dimensional supersonic turbulent flow in presence of symmetric transverse injection of round jet is simulated numerically. The simulation is based on the Favre-averaged Navier-Stokes equations coupled with Wilcox’s turbulence model. The numerical solution is performed using ENO scheme and is validated with the experimental data that include the pressure distribution on the wall in front of the jet in the plane symmetry. The numerical simulation is used to investigate in detail the flow physics for a range of the pressure ratio . The well-known primary shock formations are observed (a barrel shock, a bow shock, and the system of λ-shock waves), and the vortices are identified (horseshoe vortex, an upper vortex, two trailing vortices formed in the separation region and aft of the bow shock wave, two trailing vortices that merge together into one single rotational motion). During the experiment the presence of the new vortices near the wall behind the jet for the pressure ratio is revealed.

scholarly journals Design Optimization of a Transonic-Fan Rotor Using Numerical Computations of the Full Compressible Navier-Stokes Equations and Simplex Algorithm

2014 ◽  
Vol 2014 ◽  
pp. 1-16 ◽  
Author(s):  
M. A. Aziz ◽  
Farouk M. Owis ◽  
M. M. Abdelrahman
Keyword(s):  
Stokes Equations ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Simplex Algorithm ◽  
Navier Stokes ◽  
Numerical Computations ◽  
Navier Stokes Equations ◽  
Pressure Losses ◽  
Optimization Parameters ◽  
Transonic Fan

The design of a transonic-fan rotor is optimized using numerical computations of the full three-dimensional Navier-Stokes equations. The CFDRC-ACE multiphysics module, which is a pressure-based solver, is used for the numerical simulation. The code is coupled with simplex optimization algorithm. The optimization process is started from a suitable design point obtained using low fidelity analytical methods that is based on experimental correlations for the pressure losses and blade deviation angle. The fan blade shape is defined by its stacking line and airfoil shape which are considered the optimization parameters. The stacking line is defined by lean, sweep, and skews, while blade airfoil shape is modified considering the thickness and camber distributions. The optimization has been performed to maximize the rotor total pressure ratio while keeping the rotor efficiency and surge margin above certain required values. The results obtained are verified with the experimental data of Rotor 67. In addition, the results of the optimized fan indicate that the optimum design is found to be leaned in the direction of rotation and has a forward sweep from the hub to mean section and backward sweep to the tip. The pressure ratio increases from 1.427 to 1.627 at the design speed and mass flow rate.

scholarly journals Детально-сравнительный анализ взаимодействия сверхзвукового потока с поперечной газовой струей при больших параметрах нерасчетности

2019 ◽  
Vol 89 (10) ◽  
pp. 1513
Author(s):  
А.О. Бекетаева ◽  
P. Bruel ◽  
А.Ж. Найманова
Keyword(s):  
Gas Flow ◽  
Stokes Equations ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Mach Disk ◽  
Navier Stokes ◽  
Third Order ◽  
Navier Stokes Equations ◽  
The Third ◽  
Eno Scheme

The interaction of the spatial supersonic turbulent gas flow with a sound jet injected perpendicularly was widely studied both numerically and experimentally. However, there are only a few studies of the detail analysis of the formation and distribution of vortex structures from moderate till high pressure ratio (the ratio of pressure in the jet to pressure in the main flow).The aim of this paper is the study and identify the system of the vortex forming behind the injected sound jet in a transverse supersonic flow from the point of view of the mixing efficiency. For that the three-dimensional Favre-averaged Navier-Stokes equations, coupled with the turbulence model are solved numerically on the basis of the third-order ENO scheme. The three-dimensional Favre-averaged Navier-Stokes equations, coupled with the turbulence model are solved numerically on the basis of the third-order ENO scheme. The presence of well known vortex structures are shown: two oppositely rotating vortices in front of the jet; horseshoe vortex; two pairs of the vortex in the mixing zone of the jet and the main flow, where one of them is located in the wake behind the jet and other in the lateral line of the jet. Also, the pressure ratio parameters are determined at which the additional pairs of vortices appear. Where, the first of them is formed on the edge of the Mach disk as a result of the interaction of the decelerated jet flow behind the Mach disk with the high-speed ascending flow behind the barrel. And, the second is due to the interaction of the ascending jet flow with the main gas flow. As a result of comparative analysis the criterion of the pressure ratio parameters are found under which a clear picture of additional horn vortices is observed near the wall in the region behind the jet. The graph of the dependence of the angle of inclination of the bow shock wave on the parameter of pressure ratio is obtained. Satisfactory agreement of the pressure distribution on the wall in front of the jet in the symmetry plane with experimental data is established.

Aerodynamic Performances of a Centrifugal Compressor With Discrete Cavities

2017 ◽  
Author(s):  
Sang-Bum Ma ◽  
Kwang-Yong Kim
Keyword(s):  
Centrifugal Compressor ◽  
Stokes Equations ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Navier Stokes Equations ◽  
Stall Margin ◽  
Tetrahedral Grid ◽  
Aerodynamic Performances ◽  
Inclined Angle

In this study, aerodynamic performance of the centrifugal compressor was investigated by using a recirculating device called discrete cavities. A parametric study was conducted using six parameters related to the geometry of the discrete cavities, i.e., the inclined angle, the port angle and width, the length of cavity, the axial location of each cavity, and the number of cavities. Three-dimensional Reynolds-averaged Navier-Stokes equations with shear stress transport turbulence model were used for aerodynamic analysis of the centrifugal compressor with discrete cavities. The hexahedral grid was used in impeller domain and tetrahedral grid was used in volute and inclined discrete cavities. The numerical results for the adiabatic efficiency and the total pressure ratio (inlet to outlet) showed good agreements with experimental data. It was found that inclination of the discrete cavities further increased the stall margin without loss of efficiency compared to the uninclined discrete cavities.

Design of a Highly Loaded Transonic Two-Stage Fan Using Swept and Bowed Blading

2011 ◽  
Author(s):  
Hailiang Jin ◽  
Donghai Jin ◽  
Fang Zhu ◽  
Ke Wan ◽  
Xingmin Gui
Keyword(s):  
Total Pressure ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Leading Edge ◽  
Navier Stokes ◽  
Two Stage ◽  
Stall Margin ◽  
Stator Vane ◽  
Total Pressure Ratio ◽  
Highly Loaded

This paper presents the design of a highly loaded transonic two-stage fan using several advanced three-dimensional blading techniques including forward sweep and “hub bending” in rotors and several bowed configurations in stators. The effects of these blading techniques on the performance of the highly loaded transonic two-stage fan were investigated on the basis of three-dimensional Navier-Stokes predictions. The results indicate that forward sweep has insignificant impact on the total pressure ratio and adiabatic efficiency of the fan. The throttling range of the fan is found to be improved by forward sweep because the shock in the forward swept rotor is expelled later upstream to the leading edge than that in the unswept one. Hub bending design technique increases the efficiency in the hub region of R1 due to the reduction of the low momentum zone in the hub region near the trailing edge. The stator vane design has a pronounced impact on the performance of the fan. The total pressure ratio, adiabatic efficiency, and stall margin of the schemes with the bowed vanes are increased significantly compared to the scheme with the straight vanes. The large corner stall in the straight S1 vane is reduced effectively by the bowed S1 vanes. Moreover, the strong corner stall in the straight S2 vane is fully eliminated by the bowed S2 vanes. Among the bowed vane schemes, the scheme with positive bowed (P. B.) hub and negative bowed (N. B.) tip vanes has the best efficiency and stall margin performances thanks to the superiority of the performance over the midspan regions of the bowed vanes.

The Calculation of Three-Dimensional Viscous Flow Through Multistage Turbomachines

1992 ◽  
Vol 114 (1) ◽  
pp. 18-26 ◽  
Author(s):  
J. D. Denton
Keyword(s):  
Stokes Equations ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Viscous Effects ◽  
Navier Stokes Equations ◽  
High Pressure Ratio ◽  
Machine Performance ◽  
Multistage Turbine ◽  
Flow Through

The extension of a well-established three-dimensional flow calculation method to calculate the flow through multiple turbomachinery blade rows is described in this paper. To avoid calculating the unsteady flow, which is inherent in any machine containing both rotating and stationary blade rows, a mixing process is modeled at a calculating station between adjacent blade rows. The effects of this mixing on the flow within the blade rows may be minimized by using extrapolated boundary conditions at the mixing plane. Inviscid calculations are not realistic for multistage machines and so the method includes a range of options for the inclusion of viscous effects. At the simplest level such effects may be included by prescribing the spanwise variation of polytropic efficiency for each blade row. At the most sophisticated level viscous effects and machine performance can be predicted by using a thin shear layer approximation to the Navier–Stokes equations and an eddy viscosity turbulence model. For high-pressure-ratio compressors there is a strong tendency for the calculation to surge during the transient part of the flow. This is overcome by the use of a new technique, which enables the calculation to be run to a prescribed mass flow. Use of the method is illustrated by applying it to a multistage turbine of simple geometry, a two-stage low-speed experimental turbine, and two multistage axial compressors.

Effects of stator splitter blades on aerodynamic performance of a single-stage transonic axial compressor

2020 ◽  
Vol 14 (4) ◽  
pp. 7369-7378
Author(s):  
Ky-Quang Pham ◽  
Xuan-Truong Le ◽  
Cong-Truong Dinh
Keyword(s):  
Stokes Equations ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Aerodynamic Performance ◽  
Numerical Results ◽  
Single Stage ◽  
Navier Stokes ◽  
Navier Stokes Equations ◽  
Operating Stability ◽  
Length Coverage

Splitter blades located between stator blades in a single-stage axial compressor were proposed and investigated in this work to find their effects on aerodynamic performance and operating stability. Aerodynamic performance of the compressor was evaluated using three-dimensional Reynolds-averaged Navier-Stokes equations using the k-e turbulence model with a scalable wall function. The numerical results for the typical performance parameters without stator splitter blades were validated in comparison with experimental data. The numerical results of a parametric study using four geometric parameters (chord length, coverage angle, height and position) of the stator splitter blades showed that the operational stability of the single-stage axial compressor enhances remarkably using the stator splitter blades. The splitters were effective in suppressing flow separation in the stator domain of the compressor at near-stall condition which affects considerably the aerodynamic performance of the compressor.

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