Realistic Inlet Distortion Patterns Interacting with a Transonic Compressor Stage

2015 ◽  
pp. 285-302 ◽  
Author(s):  
Fabian Wartzek ◽  
Felix Holzinger ◽  
Christoph Brandstetter ◽  
Heinz-Peter Schiffer
Keyword(s):  
Compressor Stage ◽  
Transonic Compressor ◽  
Inlet Distortion

Effects of Inlet Distortion on the Flow Field in a Transonic Compressor Rotor

1996 ◽  
Author(s):  
Chunill Hah ◽  
Douglas C. Rabe ◽  
Thomas J. Sullivan ◽  
Aspi R. Wadia
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Viscous Flow ◽  
Total Pressure ◽  
Stokes Equations ◽  
Three Dimensional ◽  
Aerodynamic Loss ◽  
Transonic Compressor ◽  
Inlet Distortion ◽  
Compressor Rotor

The effects of circumferential distortions in inlet total pressure on the flow field in a low-aspect-ratio, high-speed, high-pressure-ratio, transonic compressor rotor are investigated in this paper. The flow field was studied experimentally and numerically with and without inlet total pressure distortion. Total pressure distortion was created by screens mounted upstream from the rotor inlet. Circumferential distortions of 8 periods per revolution were investigated at two different rotor speeds. The unsteady blade surface pressures were measured with miniature pressure transducers mounted in the blade. The flow fields with and without inlet total pressure distortion were analyzed numerically by solving steady and unsteady forms of the Reynolds-averaged Navier-Stokes equations. Steady three-dimensional viscous flow calculations were performed for the flow without inlet distortion while unsteady three-dimensional viscous flow calculations were used for the flow with inlet distortion. For the time-accurate calculation, circumferential and radial variations of the inlet total pressure were used as a time-dependent inflow boundary condition. A second-order implicit scheme was used for the time integration. The experimental measurements and the numerical analysis are highly complementary for this study because of the extreme complexity of the flow field. The current investigation shows that inlet flow distortions travel through the rotor blade passage and are convected into the following stator. At a high rotor speed where the flow is transonic, the passage shock was found to oscillate by as much as 20% of the blade chord, and very strong interactions between the unsteady passage shock and the blade boundary layer were observed. This interaction increases the effective blockage of the passage, resulting in an increased aerodynamic loss and a reduced stall margin. The strong interaction between the passage shock and the blade boundary layer increases the peak aerodynamic loss by about one percent.

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.

Optimization of Robust Transonic Compressor Blades

2016 ◽  
Author(s):  
Marcus Lejon ◽  
Niklas Andersson ◽  
Tomas Grönstedt ◽  
Lars Ellbrant ◽  
Hans Mårtensson
Keyword(s):  
Surface Roughness ◽  
Service Use ◽  
Axial Compressor ◽  
Optimization Approach ◽  
Optimized Design ◽  
Compressor Stage ◽  
Compressor Blades ◽  
Transonic Compressor ◽  
Long Period ◽  
High Level

Surface degradation in an axial compressor during its lifetime can have a considerable adverse effect on its performance. The present study investigates how the optimized design of compressor blades in a single compressor stage is affected by considering a high level of surface roughness on a level representative of a long period of in-service use. It is shown that including surface roughness in the optimization process is of relatively little importance, however, matching of compressor stages is shown to require consideration as the rotational speed must be increased to reach the design point as surface quality decrease. An increased surface roughness in itself is shown to have a large effect on performance. Two optimization approaches are compared. The first approach considers the compressor blades to be hydraulically smooth. The designs obtained from this approach are subsequently degraded by increasing the level of surface roughness. The compressor blades from the first approach are compared to designs obtained from a second optimization approach, which considers a high level of surface roughness from the outset. The degraded compressor stages from the first approach are shown to be among the best performing designs in terms of polytropic efficiency and stability when compared to designs obtained with the second approach.

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.

Scaling of Incidence Variations With Inlet Distortion for a Transonic Axial Compressor

2020 ◽  
Vol 142 (2) ◽  
Author(s):  
D. J. Hill ◽  
J. J. Defoe
Keyword(s):  
Total Pressure ◽  
Incidence Angle ◽  
Computational Cost ◽  
Force Model ◽  
Flow Angle ◽  
Total Enthalpy ◽  
Transonic Compressor ◽  
Inlet Distortion ◽  
Scaling Relationships ◽  
The Impact

Abstract This paper numerically explores the manner in which blade row inlet incidence variation scales with various distortion patterns and intensities. The objectives are to (1) identify the most appropriate parameter whose circumferential variation can be used to assess scaling relationships of a transonic compressor and (2) use this parameter to evaluate two types of non-uniform inflow patterns, vertically stratified total pressure distortions and radially stratified total enthalpy and total pressure distortions. A body force model of the blade rows is employed to reduce computational cost; the approach has been shown to capture distortion transfer and to be able to predict upstream flow redistribution with inlet distortion. Diffusion factor is shown to be an inadequate proxy for streamline loss generation in non-uniform flow, leading to the choice of incidence angle variations as the metric for which we assess scaling relationships. Posteriori scaling of circumferential flow angle variation based on the maximum incidence excursion for varying distortion intensity yields an accurate method for prediction of the impact for other distortion intensities; linear regression of the maximum variation in incidence around the annulus as a function of distortion intensity had R2 > 0.97 for all spanwise locations examined in both the rotor and stator for both vertically and radially stratified distortions. However, changes to far upstream distortion shape yield highly non-linear incidence variation scaling; the results suggest that the pitchwise gradients of far upstream total pressure govern the degree of linearity for incidence variation scaling.

A Transonic Compressor Stage: Part 2 — CFD Simulation

2004 ◽  
Author(s):  
G. V. Hobson ◽  
A. J. Gannon ◽  
R. P. Shreeve
Keyword(s):  
Cfd Simulation ◽  
Pressure Ratio ◽  
Internal Flow ◽  
Design Tool ◽  
Experimental Results ◽  
Navier Stokes ◽  
Experimental Program ◽  
Compressor Stage ◽  
Transonic Compressor ◽  
Cfd Method

The simulation of a transonic compressor stage is presented. This stage was designed using an Euler CFD code with the intent of minimizing the use of empirical design techniques. The stage has subsequently been built and tested. More recently an existing multi-block Navier-Stokes code with a steady averaging-plane to pass information between the blade rows was used to simulate the flow through the machine. Performance maps of stage pressure ratio and efficiency at 70, 80, 90 and 100% speeds from both the Euler and Navier-Stokes CFD codes are compared with the experimental results. Details of the internal flow from the Navier-Stokes code are presented. Comparison of the design Euler CFD and experimental results showed reasonable agreement and validated its use as a design tool. Agreement between experimental and the current Navier-Stokes CFD results was good, allowing the code to be used in the viewing of the internal flow field. Improvements to the initial design CFD method are discussed in light of the experimental program and more recent simulations.

Experimental Investigation on the Effect of Porosity of Bend Skewed Casing Treatment on a Single Stage Transonic Axial Flow Compressor

2014 ◽  
Author(s):  
Dilipkumar B. Alone ◽  
Subramani Satish Kumar ◽  
Shobhavathy Thimmaiah ◽  
Janaki Rami Reddy Mudipalli ◽  
A. M. Pradeep ◽  
...  
Keyword(s):  
Experimental Investigation ◽  
Axial Flow ◽  
Compressor Stage ◽  
Casing Treatment ◽  
Operating Range ◽  
Axial Flow Compressor ◽  
Transonic Compressor ◽  
Stable Operating ◽  
Flow Compressor ◽  
Casing Treatments

A bend skewed casing treatment was designed, to study the influence of one of its geometrical parameter porosity on the stable performance of single stage transonic axial flow compressor. The compressor was designed for the stage total-to-total pressure ratio of 1.35, corrected mass flow rate of 22 kg/s at corrected design speed of 12930 RPM. Bend skewed casing treatment has an axial inlet segment till 50% of the total length and rear segment that is skewed by 45° in the direction of the rotor tip section stagger. Both the sections were oriented at a skew angle of 45° to the radial plane such that the flow exiting the slot is in counter-clockwise direction to that of the rotor direction. The casing treatment slot width was equal to the maximum thickness of the rotor blades. Three casing treatment configurations were identified for the current experimental investigation. All the treatment geometries considered for the experimental research have lower porosities than reported in the open literatures. The effect of the porosity parameter on the performance of transonic compressor stage was evaluated at two axial coverages of 20% and 40% relative to the rotor tip axial chord. Performance maps were obtained for the solid casing and casing treatment with three different porosities. Comparative studies were carried out and experimental results showed a maximum of 65% improvement in the stable operating range of the compressor for one of the treatment configurations. It was also observed that the stable operating range of the compressor increases with an increase in the casing treatment porosity. All the casing treatment configurations showed that the compressor stall occurs at lower mass flows as compared to the solid casing. Compressor stage peak efficiency shows significant degradations with increase in the porosity as compared to solid casing. Detailed blade element performances were also obtained using calibrated multi-hole aerodynamic probe. Comparative variations of flow parameters like absolute flow angle, Mach number were studied at full flow and near stall conditions for the solid casing and casing treatment configurations. Hot wire measurements show very high fluctuation in the inlet axial velocity in the presence of solid casing as compared to casing treatments. Experimental investigation revealed that the porosity of the casing treatments has strong influence on the transonic compressor stage performance.

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