scholarly journals Assessment of the Severity of Unsteady Mach Number Effects in a 3-Stage Transonic Compressor

2017 ◽  
Author(s):  
Anthony Dent ◽  
Liping Xu ◽  
Roger Wells
Keyword(s):  
Leading Edge ◽  
Compressor Stage ◽  
Pressure Potential ◽  
Cfd Simulations ◽  
Transonic Compressor ◽  
Inlet Guide Vanes ◽  
Great Consideration ◽  
Stator Blade ◽  
Front Stage ◽  
Stage Efficiency

In this paper results from steady and unsteady CFD simulations of an industrial transonic compressor are compared, in order to gain a better understanding of the cause of the differences in the predicted efficiencies between the steady and unsteady simulations. Initially the first stage is simulated as an isolated compressor stage with inlet guide vanes in order to analyse the effect of individual blade rows on the stage performance. It is found that the rotor efficiency is lower for steady simulations than for unsteady simulations due to stronger shock waves. The stator efficiency is greater in the steady simulations due to not being able to model the interaction of the rotor wakes with the stator blade leading edge and boundary layers. Greater variation between steady and unsteady predictions is found at higher operating speeds. In the 3-stage unsteady simulations, the front stage efficiency characteristic is the same as the efficiency calculated from the isolated unsteady simulations. This shows that the unsteady pressure potential propagating from the downstream stages has no significant effect on the front stage efficiency meaning that the designer does not need to give great consideration to the downstream blade rows when predicting the characteristics of the front stage.

Effect of Airfoil Vortex Generator on the Performance and Stability of a Transonic Axial Compressor

2021 ◽  
Author(s):  
Subbaramu Shivaramaiah ◽  
Mahesh K. Varpe
Keyword(s):  
Pressure Ratio ◽  
Research Work ◽  
Vortex Generator ◽  
Leading Edge ◽  
Leakage Flow ◽  
Compressor Stage ◽  
Tip Leakage Flow ◽  
Stall Margin ◽  
Tip Leakage ◽  
Transonic Compressor

Abstract In the present research work, effect of airfoil vortex generator on performance and stability of transonic compressor stage is investigated through CFD simulations. In turbomachines vortex generators are used to energize boundary and generated vortex is made to interact with tip leakage flow and secondary flow vortices formed in rotor and stator blade passage. In the present numerical investigation symmetrical airfoil vortex generator is placed on rotor casing surface close to leading edge, anticipating that vortex generated will be able to disturb tip leakage flow and its interaction with rotor passage core flow. Six different vortex generator configuration are investigated by varying distance between vortex generator trailing edge and rotor leading edge. Particular vortex generator configuration shows maximum improvement of stall margin and operating range by 5.5% and 76.75% respectively. Presence of vortex generator alters flow blockage by modifying flow field in rotor tip region and hence contributes to enhancement of stall margin. As a negative effect, interaction of vortex generator vortices and casing causes surface friction and high entropy generation. As a result compressor stage pressure ratio and efficiency decreases.

Effect of Rotor Tip Casing Treatment on the Performance and Stability of a Transonic Axial Compressor

2021 ◽  
Author(s):  
Subbaramu Shivaramaiah ◽  
Mahesh K. Varpe ◽  
Mohammed Afzal
Keyword(s):  
Research Work ◽  
Stability Margin ◽  
Leakage Flow ◽  
Compressor Stage ◽  
Tip Leakage Flow ◽  
Stall Margin ◽  
Casing Treatment ◽  
Tip Leakage ◽  
Transonic Compressor ◽  
Stage Efficiency

Abstract In a transonic compressor rotor, tip leakage flow interacts with passage shock, casing boundary layer and secondary flow. This leads to increase in total pressure loss and reduction of compressor stability margin. Casing treatment is one of the passive endwall geometry modification technique to control tip leakage flow interaction. In the present investigation effect of rotor tip casing treatment is investigated on performance and stability of a NASA 37 transonic compressor stage. Existing literature reveals, that endwall casing treatment slots i.e., porous casing treatment, axial slots axially skewed slots, circumferential grooves, recirculating casing treatment etc. are able to improve compressor stability margin with penalty on stage efficiency. Turbomachinery engineers and scientists are still focusing their research work to identify an endwall casing treatment configuration with improves both compressor stall margin as well as stage efficiency. Hence in the current work, as an innovative idea, effect of casing treatment slot along rotor tip mean camber line is investigated on NASA 37 compressor stage. Casing treatment slot with rectangular cross-section was created along the rotor tip mean camber line. Four different casing treatment configurations were created by changing number of slots on rotor casing surface. In all four configurations casing treatment slot width and height remains same. Flow simulation of NASA 37 compressor stage was performed with all these four casing treatment configurations. A maximum stall margin improvement of 3% was achieved with a particular slot configuration, but without any increase in compressor stage efficiency.

An Experimental Study of Vane Clocking Effects on Embedded Compressor Stage Performance

2008 ◽  
Author(s):  
Nicole L. Key ◽  
Patrick B. Lawless ◽  
Sanford Fleeter
Keyword(s):  
Effective Means ◽  
Leading Edge ◽  
Maximum Efficiency ◽  
Frequency Noise ◽  
Compressor Stage ◽  
Performance Measurements ◽  
Refining Performance ◽  
Stage Performance ◽  
Vane Clocking ◽  
Stage Efficiency

Previous research has shown that vane clocking, the circumferential indexing of adjacent vane rows with similar vane counts, can be an effective means to increase stage performance, reduce discrete frequency noise, and/or reduce the unsteady blade forces that can lead to high cycle fatigue. The objective of this research was to experimentally investigate the effects of vane clocking in an embedded compressor stage, focusing on stage performance. Experiments were performed in the intermediate-speed Purdue 3-Stage Compressor, which consists of an IGV followed by three stages. The IGV, Stator 1, and Stator 2 vane rows have identical vane counts, and the effects of vane clocking were studied on Stage 2. Much effort went into refining performance measurements to enable the detection of small changes in stage efficiency associated with vane clocking. At design loading, the change in stage efficiency between the maximum and minimum efficiency clocking configurations was 0.27 points. The maximum efficiency clocking configuration positioned the Stator 1 wake at the Stator 2 leading edge. This condition produced a shallower and thinner Stator 2 wake compared to the clocking configuration that located the wake in the middle of the Stator 2 passage. At high loading, the change in Stage 2 efficiency associated with vane clocking effects increased to 1.07 points; however, the maximum efficiency clocking configuration was the case where the Stator 1 wake passed through the middle of the downstream vane passage. Thus, impingement of the upstream vane wake on the downstream vane leading edge resulted in the best performance at design point but provided the lowest efficiency at an off-design condition.

scholarly journals Development and Aerothermal Investigation of Integrated Combustor Vane Concept

2015 ◽  
Vol 138 (1) ◽  
Author(s):  
Simon Jacobi ◽  
Budimir Rosic
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Gas Turbines ◽  
High Speed ◽  
Leading Edge ◽  
Cfd Simulations ◽  
Pressure Loss Coefficient ◽  
Novel Concept ◽  
Nozzle Guide ◽  
Stage Efficiency

This paper presents the development and aerothermal investigation of the integrated combustor vane concept for power generation gas turbines with individual can combustors. In this novel concept, first introduced in 2010, the conventional nozzle guide vanes (NGVs) are removed and flow turning is achieved by vanes that extend the combustor walls. The concept was developed using the in-house computational fluid dyanamics (CFD) code TBLOCK. Aerothermal experiments were conducted using a modular high-speed linear cascade, designed to model the flow at the combustor–vane interface. The facility is comprised of two can combustor transition ducts and either four conventional vanes (CVs) or two integrated vanes (IVs). The experimental study validates the linear CFD simulations of the IV development. Annular full-stage CFD simulations, used to evaluate aerodynamics, heat transfer, and stage efficiency, confirm the trends of the linear numerical and experimental results, and thus demonstrate the concept's potential for real gas turbine applications. Results show a reduction of the total pressure loss coefficient at the exit of the stator vanes by more than 25% due to a reduction in profile and endwall loss. Combined with an improved rotor performance demonstrated by unsteady stage simulations, these aerodynamic benefits result in a gain in stage efficiency of above 1%. A distinct reduction in heat transfer coefficient (HTC) levels on vane surfaces, on the order of 25–50%, and endwalls is observed and attributed to an altered state of boundary layer (BL) thickness. The development of IV's endwall- and leading edge (LE)-cooling geometry shows a superior surface coverage of cooling effectiveness, and the cooling requirements for the first vane are expected to be halved. Moreover, by halving the number of vanes, simplifying the design and eliminating the need for vane LE film cooling, manufacturing and development costs can be significantly reduced.

Unsteady Forcing vs. Efficiency: The Effect of Clocking on a Transonic Industrial Compressor

2013 ◽  
Author(s):  
Florian Fruth ◽  
Damian M. Vogt ◽  
Ronnie Bladh ◽  
Torsten H. Fransson
Keyword(s):  
Stokes Equations ◽  
Leading Edge ◽  
Forced Response ◽  
Navier Stokes ◽  
Navier Stokes Equations ◽  
Transonic Compressor ◽  
Design Changes ◽  
Transient Simulations ◽  
Stator Blade ◽  
The Impact

A numerical investigation on the impact of clocking on the efficiency and the aerodynamic forcing of the first 1.5 stages of an industrial transonic compressor was conducted. Using unsteady 3D Navier-Stokes equations, seven clocking positions were calculated and analyzed. Efficiency changes due to clocking were up to 0.125%, whereas modal excitation changes up to 31.7%. However, no direct correlation between the parameters of efficiency, stimulus and modal excitation was found as reported by others. It was found that potential forced response risks can be reduced by clocking, resulting only in minor efficiency penalties. Assuming almost sinusoidal behavior of efficiency and stimulus changes, as found in this investigation, both parameters can be set into correlation by using an ellipse interpolation. Direct impact of design changes on efficiency and stimulus through clocking can be deducted from that graph and quick estimations about extrema be made using only 5–6 transient simulations. Results however also stress the importance of considering modal excitation when optimizing for aerodynamic forcing, for which the ellipse interpolation is not necessarily possible. Highest efficiency is achieved with the IGV wake impinging on the stator blade leading edge at mid-span. It was found however that this alone is not a sufficient criteria in case of inclined wakes, as wake impingement at different span positions leads to different efficiencies.

Performance Characterization of the Effect of Axial Positioning of Bend Skewed Casing Treatment Retrofitted to a Transonic Axial Flow Compressor

2014 ◽  
Author(s):  
Dilipkumar B. Alone ◽  
Subramani Satish Kumar ◽  
Shobhavathy Thimmaiah ◽  
Janaki Rami Reddy Mudipalli ◽  
A. M. Pradeep ◽  
...  
Keyword(s):  
Axial Flow ◽  
Single Stage ◽  
Compressor Stage ◽  
Stall Margin ◽  
Casing Treatment ◽  
Axial Flow Compressor ◽  
Transonic Compressor ◽  
Maximum Improvement ◽  
Flow Compressor ◽  
Stage Efficiency

The performance of an aero-engines to a large extend depends on the performance behavior of axial flow compressors and is restricted by the compressor instabilities like rotating stall and surge. In the present study, attempts have been made to design and develop the bend skewed casing treatment geometries with lower porosities to improve the stable operating range of single stage axial flow compressor. Experimental investigations were undertaken to study the impact of axial position of one of the casing treatment geometry on the single stage transonic axial flow compressor. The transonic compressor used for the current experimental studies has a stage total to total pressure ratio of 1.35, corrected mass flow rate of 22 kg/s at an operating speed of 12930 rpm. The compressor stage steady and unsteady state response for 20%, 40%, 60% and 100% axial chord coverage relative to the rotor tip chord of the bend skewed casing treatment with a porosity of 33% was studied experimentally. The objective was to identify the optimum axial location; which will give maximum improvement in the stall margin with minimal loss of compressor stage efficiency. Through an experimental study it was observed that the axial location of bend skewed casing treatment plays a very crucial role in governing the performance of the transonic compressor. For all the investigated axial coverages, compressor stall margin increases but the optimum performance in terms of stall margin improvement and efficiency gains were observed at 20% and 40% of the rotor chord. This trend shows good agreement with existing published literature. An improvement of 31.7% in the stall margin with an increase in the stage efficiency was obtained at one of the axial coverage. Maximum improvement of 37% in the stall margin above the solid casing was noticed at 60% axial coverage. The stalling characteristics of the compressor stage also changes with the axial positions. In the presence of solid casing the nature of stall was abrupt and stalls cells travels at half the rotor speed. The blade element performance also studied at the rotor exit using pre-calibrated aerodynamic probe.

Transonic radial compressor inlet design

2003 ◽  
Vol 217 (4) ◽  
pp. 367-374 ◽  
Author(s):  
A Lohmberg ◽  
M Casey ◽  
S Ammann
Keyword(s):  
Simple Model ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Leading Edge ◽  
Cfd Simulations ◽  
Radial Compressor ◽  
Transonic Compressor ◽  
Blade Thickness ◽  
Compressor Inlet ◽  
Computational Fluid Dynamics Cfd

The design of radial compressor inlets for transonic flow is examined. A theoretical model [1] quantifies the losses in the tip sections caused by the choke margin (incidence) and the blockage of the blades. It identifies clear design rules for the tip sections: to achieve the highest efficiency, these require minimum blockage (low blade thickness and splitter vanes) and low choke margin (close to the unique-incidence condition). Simulations of the NASA rotor 37 transonic axial compressor (with CFX-TASCflow) are used to validate the use of three-dimensional viscous computational fluid dynamics (CFD) for transonic compressor inlets and to demonstrate that the key performance features suggested by the simple model are also modelled in three-dimensional viscous flow simulations. The simple model together with CFD simulations has been used for the design of tip sections at the inlet of a transonic radial compressor. CFD simulations were used to select the position of the shock to give a low choke margin, to reduce the preshock Mach number and also to optimize the shape and position of the leading edge of the splitter vanes.

An Experimental Study of Vane Clocking Effects on Embedded Compressor Stage Performance

2009 ◽  
Vol 132 (1) ◽  
Author(s):  
Nicole L. Key ◽  
Patrick B. Lawless ◽  
Sanford Fleeter
Keyword(s):  
Effective Means ◽  
Leading Edge ◽  
Maximum Efficiency ◽  
Frequency Noise ◽  
Compressor Stage ◽  
Performance Measurements ◽  
Refining Performance ◽  
Stage Performance ◽  
Vane Clocking ◽  
Stage Efficiency

Previous research has shown that vane clocking, the circumferential indexing of adjacent vane rows with similar vane counts, can be an effective means to increase stage performance, reduce discrete frequency noise, and/or reduce the unsteady blade forces that can lead to high cycle fatigue. The objective of this research was to experimentally investigate the effects of vane clocking in an embedded compressor stage, focusing on stage performance. Experiments were performed in the intermediate-speed Purdue three-stage compressor, which consists of an IGV followed by three stages. The IGV, Stator 1, and Stator 2 vane rows have identical vane counts, and the effects of vane clocking were studied on Stage 2. Much effort went into refining performance measurements to enable the detection of small changes in stage efficiency associated with vane clocking. At design loading, the change in stage efficiency between the maximum and minimum efficiency clocking configurations was 0.27 points. The maximum efficiency clocking configuration positioned the Stator 1 wake at the Stator 2 leading edge. This condition produced a shallower and thinner Stator 2 wake compared with the clocking configuration that located the wake in the middle of the Stator 2 passage. At high loading, the change in Stage 2 efficiency associated with vane clocking effects increased to 1.07 points; however, the maximum efficiency clocking configuration was the case where the Stator 1 wake passed through the middle of the downstream vane passage. Thus, impingement of the upstream vane wake on the downstream vane leading edge resulted in the best performance at design point but provided the lowest efficiency at an off-design condition.

Numerical and Experimental Study of the Performance Effects of Varying Vaneless Space and Vane Solidity in Radial Inflow Turbine Stators

2008 ◽  
Author(s):  
Alister Simpson ◽  
Stephen Spence ◽  
John Watterson
Keyword(s):  
Numerical Models ◽  
Leading Edge ◽  
Experimental Program ◽  
Cfd Simulations ◽  
Static Pressure Distribution ◽  
Flow Capacity ◽  
Performance Effects ◽  
Stator Vane ◽  
Design Software ◽  
Stage Efficiency

An extensive experimental program has been carried out on a 135 mm tip diameter radial turbine using a variety of stator designs, so as to facilitate direct performance comparisons of varying stator vane solidity and the effect of varying the vane-less space. A baseline nozzle was designed using commercial blade design software, having 15 vanes and a vane trailing edge to rotor leading edge radius ratio (Rte/rle) of 1.13. Two additional sets were designed and manufactured; one set having varying vane numbers of 12, 18, 24 and 30, and a further set with Rte/rle ratios of 1.05, 1.175, 1.20 and 1.25. As part of the design process a series of CFD simulations were carried out in order to guide design iterations towards achieving a matched flow capacity for each stator. In this way the variations in measured stage efficiency could be attributed to the stator passages only, thus allowing direct comparisons to be made. Inter-stage measurements were taken to capture the static pressure distribution at rotor inlet, and these measurements then used to validate subsequent numerical models. The overall losses for different stators have been quantified, and the variations in measured and computed efficiency used to recommend optimum values of vane solidity and Rte/rle.

scholarly journals Effects of Tip Leakage Flow on the Aerodynamic Performance and Stability of an Axial-Flow Transonic Compressor Stage

Energies ◽  
2021 ◽  
Vol 14 (14) ◽  
pp. 4168
Author(s):  
Botao Zhang ◽  
Xiaochen Mao ◽  
Xiaoxiong Wu ◽  
Bo Liu
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Axial Flow ◽  
Leading Edge ◽  
Rotating Stall ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Tip Leakage ◽  
Transonic Compressor

To explain the effect of tip leakage flow on the performance of an axial-flow transonic compressor, the compressors with different rotor tip clearances were studied numerically. The results show that as the rotor tip clearance increases, the leakage flow intensity is increased, the shock wave position is moved backward, and the interaction between the tip leakage vortex and shock wave is intensified, while that between the boundary layer and shock wave is weakened. Most of all, the stall mechanisms of the compressors with varying rotor tip clearances are different. The clearance leakage flow is the main cause of the rotating stall under large rotor tip clearance. However, the stall form for the compressor with half of the designed tip clearance is caused by the joint action of the rotor tip stall caused by the leakage flow spillage at the blade leading edge and the whole blade span stall caused by the separation of the boundary layer of the rotor and the stator passage. Within the investigated varied range, when the rotor tip clearance size is half of the design, the compressor performance is improved best, and the peak efficiency and stall margin are increased by 0.2% and 3.5%, respectively.

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