Transonic Axial Compressor Stage Research Articles

Shock Dynamics and Associated Flow Physics at Near Stall Condition in a Transonic Axial Flow Compressor Stage

Shock Dynamics and Associated Flow Physics at Near Stall Condition in a Transonic Axial Flow Compressor Stage

Evolution of unsteady secondary flow structures near the onset of stall in a tip-critical transonic axial flow compressor stage

The aerodynamic stability of an axial compressor stage depends on the rotor and stator. However, due to the specific design requirements, the evolving flow field and the resulting secondary flow structures are different in both components. This study investigates the evolution of dominant secondary flow structures occurring in the rotor and stator of a tip-critical transonic compressor stage at the near-stall condition using unsteady numerical analysis and validates performance characteristics with the experimental data. The investigation reveals that the presence of rotor tip shock creates a large difference in the pressure gradient across the pressure surface and the suction surface, intensifying tip leakage flow and shock-induced boundary layer separation. The higher incidence angle near the hub leading edge creates a local separation and reattachment zone. The radial pressure gradient causes the low momentum flow from this local separation zone to migrate radially upward. This migrated flow interacts with the tip leakage flow and separated blade boundary layer, eventually creating a colossal recirculation zone and subsequent rotor blockage of around 46%. The increasing streamwise adverse pressure gradient pushes the tip shock upstream, and at the onset of the stall, the flow directly separates from the rotor leading edge avoiding the shock formation. The stator flow field is dominated by the asymmetric hub corner separation induced by the streamwise adverse pressure gradient and the tip corner separation caused by the vortex structures convected downstream from the rotor tip region leading to stator blockage of around 48%. Along with the blade passing frequency, four other dominating frequencies (0.06×BPF, 0.12×BPF, 0.44×BPF, and 0.84×BPF) related to the aerodynamic instabilities are observed at the inception of stall.

Aerodynamic Performance and Stability of a Transonic Axial Compressor Stage with an Airfoil Vortex Generator

Aerodynamic Performance and Stability of a Transonic Axial Compressor Stage with an Airfoil Vortex Generator

Influence of an upstream transonic axial compressor stage on the performance of inter-stage duct

Influence of an upstream transonic axial compressor stage on the performance of inter-stage duct

Aeroelastic Aspects of Axial Compressor Stage With Self-Recirculating Casing Treatment

Abstract For successful implementation of casing treatment designs in axial compressors, apart from the stall margin improvement benefits, aeroelasticity also plays a major role. This manuscript addresses the not often discussed aeroelastic aspects of a new discrete type of passive self-recirculating casing treatment (RCT) designed for a transonic axial compressor stage. Experiments are carefully designed for synchronized measurement of the unsteady fluidic disturbances and vibrations during rotating stall for compressor with baseline solid casing and self-RCT. The modal characteristics of the axial compressor rotor-disk assembly are studied experimentally and numerically. Experimentally it is observed that the rotating stall cells excite the blades in their fundamental mode in a compressor with baseline solid casing at the stall flow condition. In contrast, there is no excitation of the blades in the compressor with self-RCT at the same solid casing stall flow condition. Also, the self-RCT compared to the solid casing can significantly reduce the overall vibration levels of the blades that are excited at the stall flow condition. The casing treatment is able to alter the flow field near the tip region of the rotor blade, and hence influencing the forcing function of the rotating cantilever blades to have the aeroelastic benefit.

Influence of Hub Contouring on the Performance of a Transonic Axial Compressor Stage with Low Hub-Tip Ratio

Influence of Hub Contouring on the Performance of a Transonic Axial Compressor Stage with Low Hub-Tip Ratio

Aerodynamic behavior of a transonic axial flow compressor stage with self-recirculating casing treatment

In this manuscript, a new discrete type of passive Self-Recirculating Casing Treatment (RCT) is designed for a transonic axial compressor stage, fabricated, and tested for its performance. Parametric evaluation of the casing flush-mounted RCT for injection skew angle is studied experimentally and numerically to achieve considerable stall margin improvements. Detailed flow investigations are carried out for the RCT configurations and baseline solid casing behavior at different flow conditions using steady and unsteady measurements. The unsteadiness in the flow and the rotating stall signatures are captured through Kulite pressure sensors. Results indicate improvements in the stall margin and other overall compressor performance parameters for the investigated speeds. Improvements of stall margin in axial compressor stage with self-recirculating casing treatment are found to be because of the RCT by investing the already worked high-pressure fluid is able to overcome losses associated with the tip leakage flow. There is a delay in the onset of rotating stall instability with casing treatment which plays a role in the enhanced performance improvements.

Parametric investigation on a transonic axial compressor stage using CFD techniques

Tip clearance and secondary flow are regarded as important parameters resulting in degradation of compressor performance as a result of the stall-inception, total loss of pressure and end-wall blockage. To investigate the effect of rotor tip clearance and rotor blade solidity on fluid dynamic interaction of tip leakage flow and secondary flow an numerical investigation was conducted on stage NASA-37 with zero rotor tip clearance, nominal and twice the nominal rotor tip clearance and also with (1.4276, 1.3574, 1.2873, 1.2145, 1.1437) rotor blade solidity. NASA-37 stage was discreatized and analysed using ANSYS-Turbogrid and CFX 15.0 software. The tip clearance flow arising from the leading edge of rotor blade flows across the passage shock and on impinging on the pressure surface of the adjacent blade interacts with vortices of secondary flow. The compressor stage performance is affected by decrease in peak stage pressure ratio by 3.6% and 2% in peak isentropic efficiency with increase in rotor tip clearance due to the strong interaction of fluid flow as a result of the retarded momentum of fluid zone. Also Low fluid flow interaction is noticed as the rotor blade solidity is increased from 1.1437 to 1.4276. Hence, reducing the losses in compressor stage at higher solidity for three tip clearances.

Numerical Investigations of Stall Development in a Transonic Axial Compressor Stage

In this study, stall development on a previously tested axial compressor stage called NASA stage 35 was investigated by computational fluid dynamics software for the purpose of overcoming computational difficulties resulting from the secondary flow and tip vortex, and the results were compared to experimental data and other relevant studies in the literature. In the computational model, the k-epsilon turbulence model and the high-resolution scheme were employed to cope with flow uncertainties and to spatial discretization of the nonlinear equations, respectively. The pressure and temperature distributions obtained from k-epsilon turbulence model were in good agreement with experimental results. However, the performance map of the compressor showed that the error values between predicted total pressure ratios and experimental results increased as the compressor reached stall due to shockwave unsteadiness.

Aerodynamic characterization of a transonic axial flow compressor stage – with asymmetric tip clearance effects

This manuscript discusses the performance of a single-stage transonic axial flow compressor with non-uniform/asymmetric rotor tip clearance. Detailed steady and unsteady experimental measurements are carried out for a compressor with non-uniform clearance over the rotor. Measurements are focused at the peak clearance region of the compressor annulus where stall inception is likely to occur. A sector-based steady-state CFD simulation is performed on the compressor stage with uniform averaged rotor running clearance obtained from the blade growth estimates at design speed using structural analysis. The role of the tip leakage vortex on the stall dynamics of the compressor is elucidated. The stall events/disturbances occurring at every two rotor revolutions are shown using hotwire measurements. Varied flow features of the transonic compressor stage are discussed for different compressor speeds investigated. The level of clearance eccentricity/asymmetry existing in this compressor due to its complex spread does not degrade the overall compressor performance behavior.

Full Annulus Simulations of a Transonic Axial Compressor Stage with Distorted Inflow at Transonic and Subsonic Blade Tip Speed

This article reports on systematic numerical studies on an axial compressor stage with distorted inflow. Four operating points at two speedlines were simulated with an inflow distortion generated by a 120 ∘ -sector segment with a wedge-type cross-section. With this setup, the interaction between the distorted inflow and the rotor flow was studied. The focus was put on the differences of the interaction between the distorted inflow and rotor flow as well as on the compressor behaviour at subsonic and transonic blade tip speeds, as the general mechanisms have been analysed in more detail in previous publications. The distorted flow itself is not influenced by the blade tip speed, but the interaction phenomena depend strongly on the spool speed and operating point. Additionally, the blade tip speed influences the circumferential sector of the compressor stage exit, which is affected by the distorted flow. The impact reaches from a small sector at 65% spool speed, with the peak efficiency operating point up to nearly the entire annulus at 100% spool speed, near the stall operating point.

Effect of Rotor Tip Casing Treatment on the Performance and Stall Margin of a Transonic Axial Compressor Stage

Effect of Rotor Tip Casing Treatment on the Performance and Stall Margin of a Transonic Axial Compressor Stage

An Innovative Method for the Evaluation of Particle Deposition Accounting for Rotor/Stator Interaction

Solid particle ingestion is one of the principal degradation mechanisms in the compressor and turbine sections of gas turbines. In particular, in industrial applications, the microparticles not captured by the air filtration system can cause deposits on blading and, consequently, result in a decrease in the compressor performance. This paper presents three-dimensional numerical simulations of the microparticle ingestion (0.15–1.50 μm) in a transonic axial compressor stage, carried out by means of a commercial computational fluid dynamic code. Particles of this size can follow the main air flow with relatively little slip, while being impacted by the flow turbulence. It is of great interest to the industry to determine which zones of the compressor blades are impacted by these small particles. Particle trajectory simulations use a stochastic Lagrangian tracking method that solves the equations of motion separately from the continuous phase. A particular computational strategy is adopted in order to take into account the presence of two subsequent annular cascades (rotor and stator) in the case of particle ingestion. The proposed strategy allows the evaluation of particle deposition in an axial compressor stage, thanks to its capability of accounting for rotor/stator interaction. NASA Stage 37 is used as a case study for the numerical investigation. The compressor stage numerical model and the discrete phase model are set up and validated against the experimental and numerical data available in the literature. The blade zones affected by the particle impact and the kinematic characteristics of the impact of micrometric and submicrometric particles with the blade surface are shown. Both blade zones affected by the particle impact and deposition are analyzed. The particle deposition is established by using the quantity called sticking probability, adopted from the literature. The sticking probability links the kinematic characteristics of particle impact on the blade with fouling phenomenon. The results show that microparticles tend to follow the flow by impacting at full span with a higher impact concentration on the pressure side of rotor blade and stator vane. Both the rotor blade and stator vane suction side are only affected by the impact of smaller particles (up to 1 μm). Particular fluid dynamic phenomena, such as separation, shock waves, and tip leakage vortex, strongly influence the impact location of the particles. The kinematic analysis shows a high tendency of particle adhesion on the suction side of the rotor blade, especially for particles with a diameter equal to 0.15 μm.

Multi-disciplinary design optimization and performance evaluation of a single stage transonic axial compressor

The multidisciplinary design optimization method, which integrates aerodynamic performance and structural stability, was utilized in the development of a single-stage transonic axial compressor. An approximation model was created using artificial neural network for global optimization within given ranges of variables and several design constraints. The genetic algorithm was used for the exploration of the Pareto front to find the maximum objective function value. The final design was chosen after a second stage gradient-based optimization process to improve the accuracy of the optimization. To validate the design procedure, numerical simulations and compressor tests were carried out to evaluate the aerodynamic performance and safety factor of the optimized compressor. Comparison between numerical optimal results and experimental data are well matched. The optimum shape of the compressor blade is obtained and compared to the baseline design. The proposed optimization framework improves the aerodynamic efficiency and the safety factor.

다단 천음속 압축기의 유동 불안정성에 관한 실험적 연구

This study presents unsteady and unstable characteristics of three stage transonic axial compressor, developed by Korea Aerospace Research Institute. As approaching to the unstable operating region at the 103% design speed of the compressor, a modal type stall precursor appears in front of highly loaded 3rd rotor row at first, and it propagates to the upstream. On the contrary, actual stall cell initiates from the stall precursor in front of the 1st rotor row, and it propagates to the downstream of the compressor. After the stall region reached the 3rd stage and stall cell rotates circumferentially about 360 deg, it develops to one dimensional compressor surge mode. It shows a mild surge behaviour with 3∼4 Hz frequency. From the test data, it can be suggested that there is a priority to give an optimum blade loading distributions to construct a multi stage transonic axial compressor stages either to secure more stable compressor operating ranges, or to maximize the compressor efficiency.

다단 천음속 축류형 압축기 성능에 관한 실험적 연구

ABSTRACT Korea Aerospace Research Institute is performing 3 stage transonic axial compressor development program. This paper introduces design step of the compressor, the performance test results and its analysis. In the fore part of the paper, aerodynamic process of the 3 stage axial compressor is presented. To satisfy both of the mass flow and pressure rise, the compressor should rotate at a high rotational speed. Therefore the transonic flow field forms in the rotor stages and it is designed with a relatively high pressure rise per stage to satisfy its design target. The compressor stage consists of 3 stages, and the bulk pressure ratio is 2.5. The first stage is burdened with the highest pressure ratio and less pressure rises occur in the following stages. Also it is designed that tip Mach number of the first rotor row does not exceed 1.3, while the maximum relative Mach number in the rotor stage is between 1.3∼1.4 to increase the compressor flow coefficient. The final design has been confirmed by iterating three dimensional CFD calculations to verify design target and some design intentions. In the latter part of the paper, its performance test processes and results are presented. The performance test result shows that the overall compressor performance targets; pressure ratio and efficiency are well achieved. The stator static pressure distributions show that the blade loading is gradually increasing from the downstream of the compressor.

Hybrid Reynolds-Averaged Navier-Stokes/Kinetic-Eddy Simulation of Stall Inception in Axial Compressors

A hybrid Reynolds-averaged Navier-Stokes/kinetic-eddy simulation turbulence model is used for the stall predictions in a transonic axial compressor stage. This hybrid Reynolds-averaged Navier-Stokes/kinetic-eddy simulation model solves Menter's k-ω-shear-stress-transport model near walls and switches to the kinetic-eddy simulation model away from walls. The kinetic-eddy simulation model solves directly for local turbulent kinetic energy and local turbulent length scales, thus alleviating the grid spacing dependency found in other detached-eddy simulation and hybrid Reynolds-averaged Navier-Stokes/large-eddy simulation models. The current methodology is used in the prediction of the performance map and the stall inception for the NASA stage 35 compressor configuration as a representative of a modern compressor stage. The present approach is found to satisfactorily predict the onset of stall. It is found that the rotor blade-tip leakage vortex and its interaction with the shock wave is the main reason behind the stall inception in this compressor stage.

Investigation of Blade Tip Interaction With Casing Treatment in a Transonic Compressor—Part II: Numerical Results

A single stage transonic axial compressor was equipped with a casing treatment consisting of 3.5 axial slots per rotor pitch in order to investigate its influence on stall margin characteristics, as well as on the rotor near tip flow field, both numerically and experimentally. Contrary to most other studies, a generic casing treatment (CT) was designed to provide optimal optical access in the immediate vicinity of the CT, rather than for maximum benefit in terms of stall margin extension. The second part of this two-part paper deals with the numerical developments and their validation, carried out in order to efficiently perform time-accurate casing treatment simulations. The numerical developments focus on the extension of an existing coupling algorithm in order to carry out unsteady calculations with any exterior geometry coupled to the main flow passage (in this case a single slot), having an arbitrary pitch. This extension is done by incorporating frequency domain, phase-lagged boundary conditions into this coupling procedure. Whereas the phase lag approach itself is well established and validated for standard rotor-stator calculations, its application to casing treatment simulations is new. Its capabilities and validation will be demonstrated on the given compressor configuration, making extensive use of the detailed particle image velocimetry flow field measurements near the rotor tip. Instantaneous data at all measurement planes will be compared for different rotor positions with respect to the stationary slots in order to evaluate the time-dependent interaction between the rotor and the casing treatment.

Prestall Behavior of a Transonic Axial Compressor Stage via Time-Accurate Numerical Simulation

Computational fluid dynamics calculations using high-performance parallel computing were conducted to simulate the prestall flow of a transonic compressor stage, NASA compressor Stage 35. The simulations were run with a full-annulus grid that models the 3D, viscous, unsteady blade row interaction without the need for an artificial inlet distortion to induce stall. The simulation demonstrates the development of the rotating stall from the growth of instabilities. Pressure rise performance and pressure traces are compared with published experimental data before the study of flow evolution prior to the rotating stall. Spatial fast Fourier transform analysis of the flow indicates a rotating long-length disturbance of one rotor circumference, which is followed by a spike-type breakdown. The analysis also links the long-length wave disturbance with the initiation of the spike inception. The spike instabilities occur when the trajectory of the tip clearance flow becomes perpendicular to the axial direction. When approaching stall, the passage shock changes from a single oblique shock to a dual shock, which distorts the perpendicular trajectory of the tip clearance vortex but shows no evidence of flow separation that may contribute to stall.

Experimental Study of a High-Throughflow Transonic Axial Compressor Stage

Design information and experimental results are presented for a transonic axial compressor stage passing 40 lbs/s-ft2 frontal area (195 Kg/s-m2) with a pressure ratio of 1.95 at 1500 ft/s (457 m/s) tip speed. The design incorporates several unusual features that helped it achieve a peak isentropic efficiency over 88 percent at design speed. The compressor was evaluated at three rotor tip clearances and an optimum was found. Vortex generators placed upstream on the casing proved relatively ineffective in influencing stall margin. Vortex generators installed on the rotor did improve stall margin and also increased efficiency at speeds of 90 percent and below.