scholarly journals Study and Control of a Radial Vaned Diffuser Stall

2012 ◽  
Vol 2012 ◽  
pp. 1-12 ◽  
Author(s):  
Aurélien Marsan ◽  
Isabelle Trébinjac ◽  
Sylvain Coste ◽  
Gilles Leroy
Keyword(s):  
Boundary Layer ◽  
Skin Friction ◽  
Suction Side ◽  
Boundary Layer Separation ◽  
Point Of View ◽  
Vaned Diffuser ◽  
Operating Range ◽  
Stable Operating ◽  
And Control ◽  
Friction Line

The aim of the present study is to evaluate the efficiency of a boundary layer suction technique in case of a centrifugal compressor stage in order to extend its stable operating range. First, an analysis of the flow pattern within the radial vaned diffuser is presented. It highlights the stall of the diffuser vanes when reaching a low massflow. A boundary layer separation in the hub-suction side corner grows when decreasing the massflow from the nominal operating point to the surge and finally leads to a massive stall. An aspiration strategy is investigated in order to control the stall. The suction slot is put in the vicinity of the saddle that originates the main separating skin-friction line, identified thanks to the analysis of the skin-friction pattern. Several aspiration massflow rates are tested, and two different modelings of the aspiration are evaluated. Finally, an efficient control is reached with a removal of only 0,1% of the global massflow and leads—from a steady-state calculations point of view—to an increase by 40% of the compressor operating range extent.

Numerical Investigation of the Flow in a Radial Vaned Diffuser Without and With Aspiration

2012 ◽  
Author(s):  
A. Marsan ◽  
I. Trébinjac ◽  
S. Coste ◽  
G. Leroy
Keyword(s):  
Skin Friction ◽  
Leading Edge ◽  
Major Change ◽  
Suction Side ◽  
Tip Clearance ◽  
Operating Point ◽  
Vaned Diffuser ◽  
Leading Role ◽  
Stable Operating ◽  
Blade Tip

An analysis of the flow in a centrifugal compressor vaned diffuser from the nominal operating point of the compressor stage to a point near surge was conducted. Study of performance coefficients, and then of the skin-friction pattern, reveals the growth of a corner stall between the hub wall and the suction side of the vane when moving the operating point towards surge. Considering the location of the skin-friction pattern singular elements, a boundary layer suction technique has been developed and then numerically tested. The hub wall corner stall was controlled, and performances predicted near surge have been significantly modified, as well as the flow structures when reaching the limit of numerical stable operating range: the major change in the topology of the flow occurs now within the impeller, in the splitter leading edge region, and let think about a leading role of the main blade tip clearance vortex in the instabilities release. The surge massflow seems to have been significantly reduced.

An Investigation of the Stability Enhancement of a Centrifugal Compressor Stage Using a Porous Throat Diffuser

2017 ◽  
Vol 140 (1) ◽  
Author(s):  
Lee Galloway ◽  
Stephen Spence ◽  
Sung In Kim ◽  
Daniel Rusch ◽  
Klemens Vogel ◽  
...  
Keyword(s):  
Centrifugal Compressor ◽  
Compressor Stage ◽  
Vaned Diffuser ◽  
Flow Rates ◽  
Stall Inception ◽  
Centrifugal Compressor Stage ◽  
Operating Range ◽  
Stable Operating ◽  
The Stability ◽  
Circumferential Pressure

The stable operating range of a centrifugal compressor stage of an engine turbocharger is limited at low mass flow rates by aerodynamic instabilities which can lead to the onset of rotating stall or surge. There have been many techniques employed to increase the stable operating range of centrifugal compressor stages. The literature demonstrates that there are various possibilities for adding special treatments to the nominal diffuser vane geometry, or including injection or bleed flows to modify the diffuser flow field in order to influence diffuser stability. One such treatment is the porous throat diffuser (PTD). Although the benefits of this technique have been proven in the existing literature, a comprehensive understanding of how this technique operates is not yet available. This paper uses experimental measurements from a high pressure ratio (PR) compressor stage to acquire a sound understanding of the flow features within the vaned diffuser which affect the stability of the overall compression system and investigate the stabilizing mechanism of the porous throat diffuser. The nonuniform circumferential pressure imposed by the asymmetric volute is experimentally and numerically examined to understand if this provides a preferential location for stall inception in the diffuser. The following hypothesis is confirmed: linking of the diffuser throats via the side cavity equalizes the diffuser throat pressure, thus creating a more homogeneous circumferential pressure distribution, which delays stall inception to lower flow rates. The results of the porous throat diffuser configuration are compared to a standard vaned diffuser compressor stage in terms of overall compressor performance parameters, circumferential pressure nonuniformity at various locations through the compressor stage and diffuser subcomponent analysis. The diffuser inlet region was found to be the element most influenced by the porous throat diffuser, and the stability limit is mainly governed by this element.

Numerical Study of Wall Influence on Boundary-Layer Transition for Two-Dimensional NACA 16-012 and 4412 Hydrofoil Sections

1990 ◽  
Vol 34 (01) ◽  
pp. 38-47
Author(s):  
R. Latorre ◽  
R. Baubeau
Keyword(s):  
Boundary Layer ◽  
Test Section ◽  
Numerical Study ◽  
Suction Side ◽  
Boundary Layer Separation ◽  
Boundary Layer Transition ◽  
Layer Transition ◽  
Turbulent Separation ◽  
Side Pressure ◽  
Effective Angle

One of the difficulties in hydrofoil model tests is the relatively low Reynolds number of the test piece and the presence of the test section walls. This paper presents the results of systematic calculations of the potential flow field of NA 4412 and NACA 16-012 hydrofoil in a test section with wall-to-chord ratios h/c -1.0. The corresponding boundary-layer calculations using the CERT calculation scheme are presented to show the influence of the nearby walls on shifting the location of the boundary-layer laminar-turbulent separation as well as turbulent separation. By introducing an effective angle of attack, it is possible to obtain close agreement in the calculated and measured suction side pressure distortion as well as the locations of the boundary-layer separation and transition.

Unsteady Operation of a Highly Supersonic Organic Rankine Cycle Turbine

2016 ◽  
Vol 138 (12) ◽  
Author(s):  
Enrico Rinaldi ◽  
Rene Pecnik ◽  
Piero Colonna
Keyword(s):  
Boundary Layer ◽  
Organic Rankine Cycle ◽  
Suction Side ◽  
Rankine Cycle ◽  
Boundary Layer Separation ◽  
Expansion Ratio ◽  
Ideal Gas ◽  
Rotor Blades ◽  
Navier Stokes ◽  
Shock Interactions

Organic Rankine cycle (ORC) turbogenerators are the most viable option to convert sustainable energy sources in the low-to-medium power output range (from tens of kWe to several MWe). The design of efficient ORC turbines is particularly challenging due to their inherent unsteady nature (high expansion ratios and low speed of sound of organic compounds) and to the fact that the expansion encompasses thermodynamic states in the dense vapor region, where the ideal gas assumption does not hold. This work investigates the unsteady nonideal fluid dynamics and performance of a high expansion ratio ORC turbine by means of detailed Reynolds-averaged Navier–Stokes (RANS) simulations. The complex shock interactions resulting from the supersonic flow (M ≈ 2.8 at the vanes exit) are related to the blade loading, which can fluctuate up to 60% of the time-averaged value. A detailed loss analysis shows that shock-induced boundary layer separation on the suction side of the rotor blades is responsible for most of the losses in the rotor, and that further significant contributions are given by the boundary layer in the diverging part of the stator and by trailing edge losses. Efficiency loss due to unsteady interactions is quantified in 1.4% in absolute percentage points at design rotational speed. Thermophysical properties are found to feature large variations due to temperature even after the strong expansion in the nozzle vanes, thus supporting the use of accurate fluid models in the whole turbine stage.

Experimental investigations on characteristics of boundary layer and control of transition on an airfoil by AC-DBD

2018 ◽  
Vol 32 (08) ◽  
pp. 1850108 ◽  
Author(s):  
Xi Geng ◽  
Zhiwei Shi ◽  
Keming Cheng ◽  
Hao Dong ◽  
Qun Zhao ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Skin Friction ◽  
Boundary Layer Transition ◽  
Plasma Actuators ◽  
Experimental Investigations ◽  
Dbd Plasma ◽  
Layer Transition ◽  
Piv Measurement ◽  
Induced Flow ◽  
And Control

Plasma-based flow control is one of the most promising techniques for aerodynamic problems, such as delaying the boundary layer transition. The boundary layer’s characteristics induced by AC-DBD plasma actuators and applied by the actuators to delay the boundary layer transition on airfoil at Ma = 0.3 were experimentally investigated. The PIV measurement was used to study the boundary layer’s characteristics induced by the plasma actuators. The measurement plane, which was parallel to the surface of the actuators and 1 mm above the surface, was involved in the test, including the perpendicular plane. The instantaneous results showed that the induced flow field consisted of many small size unsteady vortices which were eliminated by the time average. The subsequent oil-film interferometry skin friction measurement was conducted on a NASA SC(2)-0712 airfoil at Ma = 0.3. The coefficient of skin friction demonstrates that the plasma actuators successfully delay the boundary layer transition and the efficiency is better at higher driven voltage.

Unsteady Simulation of an Axial Compressor Stage With Casing and Blade Passive Treatments

2009 ◽  
Vol 131 (2) ◽  
Author(s):  
Nicolas Gourdain ◽  
Francis Leboeuf
Keyword(s):  
Boundary Layer ◽  
Passive Control ◽  
Axial Compressor ◽  
Suction Side ◽  
Boundary Layer Separation ◽  
Leakage Flow ◽  
Compressor Stage ◽  
Tip Leakage Flow ◽  
Tip Leakage ◽  
Good Ability

This paper deals with the numerical simulation of technologies to increase the compressor performances. The objective is to extend the stable operating range of an axial compressor stage using passive control devices located in the tip region. First, the behavior of the tip leakage flow is investigated in the compressor without control. The simulation shows an increase in the interaction between the tip leakage flow and the main flow when the mass flow is reduced, a phenomenon responsible for the development of a large flow blockage region at the rotor leading edge. A separation of the rotor suction side boundary layer is also observed at near stall conditions. Then, two approaches are tested in order to control these flows in the tip region. The first one is a casing treatment with nonaxisymmetric slots. The method showed a good ability to control the tip leakage flow but failed to reduce the boundary layer separation on the suction side. However, an increase in the operability was observed but with a penalty for the efficiency. The second approach is a blade treatment that consists of a longitudinal groove built in the tip of each rotor blade. The simulation pointed out that the device is able to control partially all the critical flows with no penalty for the efficiency. Finally, some recommendations for the design of passive treatments are presented.

scholarly journals Numerical Simulation of Stall Inception Mechanisms in a Centrifugal Compressor With Vaned Diffuser

2016 ◽  
Vol 138 (12) ◽  
Author(s):  
Y. Bousquet ◽  
N. Binder ◽  
G. Dufour ◽  
X. Carbonneau ◽  
M. Roumeas ◽  
...  
Keyword(s):  
Mass Flow ◽  
Centrifugal Compressor ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Suction Side ◽  
Boundary Layer Separation ◽  
Flow Topology ◽  
Vaned Diffuser ◽  
Stall Inception ◽  
Flow Reduction

The present paper numerically investigates the stall inception mechanisms in a centrifugal compressor stage composed of a splittered unshrouded impeller and a vaned diffuser. Unsteady numerical simulations have been conducted on a calculation domain comprising all the blade passages over 360 deg for the impeller and the diffuser. Three stable operating points are simulated along a speed line, and the full path to instability is investigated. The paper focusses first on the effects of the mass flow reduction on the flow topology at the inlet of both components. Then, a detailed analysis of stall inception mechanisms is proposed. It is shown that at the inlet of both components, the mass flow reduction induces boundary layer separation on the blade suction side, which results in a vortex tube having its upper end at the casing and its lower end at the blade wall. Some similarities with flows in axial compressor operating at stall condition are outlined. The stall inception process starts with the growth of the amplitude of a modal wave rotating in the vaneless space. As the flow in the compressor is subsonic, the wave propagates upstream and interacts with the impeller flow structure. This interaction leads to the drop in the impeller pressure ratio.

Prospects of Studies in Region of Low-Sized Aircraft (Review)

2014 ◽  
Vol 9 (2) ◽  
pp. 95-115
Author(s):  
Ilya Zverkov ◽  
Alexey Kryukov ◽  
Genrich Grek
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Boundary Layer Separation ◽  
Aerodynamic Characteristics ◽  
Point Of View ◽  
Wavy Surface ◽  
Local Boundary ◽  
Effective Operation ◽  
Forward Edge ◽  
The Given

In the given review the problem of improvement of aerodynamic characteristics of the low-sized aircraft is considered with point of view of the fundamental phenomena of the mechanics of liquid, gas and plasma. It is a problem of the local boundary layer separation (separated bubbles) and flow separation from a wing forward edge at which all global structure of a flow varies. The review of the works establishing this interrelation and methods of the influence, eliminating harmful consequences of the separations is submitted. The method of separation elimination with help of a wavy surface, as the most perspective and easily sold on practice is in more details allocated in this review. The second part of the review is devoted to the analysis of a flow of elements of designs of various low-sized aircraft with indication of probably problem places where the flow is realized at Reynolds number less than 106 and where can arise the local separations. Application of a wavy surface in such places can improve aerodynamic characteristics of the flying device promoting its more effective operation

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