Free-Form Versus Ruled Inducer Design in a Transonic Centrifugal Impeller

2017 ◽  
Author(s):  
Hamid Hazby ◽  
Chris Robinson ◽  
Michael Casey ◽  
Daniel Rusch ◽  
Rene Hunziker
Keyword(s):  
Pressure Ratio ◽  
Leading Edge ◽  
Free Form ◽  
Centrifugal Impeller ◽  
Vaned Diffuser ◽  
Design Concepts ◽  
High Pressure Ratio ◽  
Mechanical Constraints ◽  
Compressor Impeller ◽  
Form Design

The detailed design of the inducer of a high pressure ratio transonic radial compressor impeller with a design inlet tip relative Mach number of 1.4 is considered. Numerical analysis has been used to compare a datum impeller with ruled inducer design with a number of different free-form design concepts, generated following the same aerodynamic design philosophy. The datum stage and one with a free-form inducer, referred to as ‘barrelled forward swept’, with forward swept leading edge near the tip and increased chord at mid-span, have been manufactured and tested. The tests were performed with the same stationary components, including the casing, vaned diffuser and the volute. The design with a barrelled forward sweep of the inducer allows the designer more control of the strength and position of the passage shock at the inlet while meeting mechanical constraints. Interestingly, the performance is also enhanced at off-design points at lower tip-speeds. The measurements show that the stage tested with the swept impeller achieves higher efficiency of between 0.5% and 1.6% compared to the datum design, depending on the operating speed. The CFD simulations are used to further study the flow at part speeds, in order to explain the causes of the observed performance differences at off design conditions.

Free-Form Versus Ruled Inducer Design in a Transonic Centrifugal Impeller

2017 ◽  
Vol 140 (1) ◽  
Author(s):  
Hamid Hazby ◽  
Chris Robinson ◽  
Michael Casey ◽  
Daniel Rusch ◽  
Rene Hunziker
Keyword(s):  
Pressure Ratio ◽  
Leading Edge ◽  
Free Form ◽  
Centrifugal Impeller ◽  
Vaned Diffuser ◽  
Design Concepts ◽  
High Pressure Ratio ◽  
Mechanical Constraints ◽  
Compressor Impeller ◽  
Form Design

The detailed design of the inducer of a high pressure ratio transonic radial compressor impeller with a design inlet tip relative Mach number of 1.4 is considered. Numerical analysis has been used to compare a datum impeller with ruled inducer design with a number of different free-form design concepts, generated following the same aerodynamic design philosophy. The datum stage and one with a free-form inducer, referred to as “barrelled forward swept,” with forward swept leading edge near the tip and increased chord at midspan, have been manufactured and tested. The tests were performed with the same stationary components, including the casing, vaned diffuser, and the volute. The design with a barrelled forward sweep of the inducer allows the designer more control of the strength and position of the passage shock at the inlet while meeting mechanical constraints. Interestingly, the performance is also enhanced at off-design points at lower tip-speeds. The measurements show that the stage tested with the swept impeller achieves higher efficiency of between 0.5% and 1.6% compared to the datum design, depending on the operating speed. The computational fluid dynamics (CFD) simulations are used to further study the flow at part speeds, in order to explain the causes of the observed performance differences at off design conditions.

Investigations of the Flow Through a High Pressure Ratio Centrifugal Impeller

2002 ◽  
Author(s):  
Gernot Eisenlohr ◽  
Hartmut Krain ◽  
Franz-Arno Richter ◽  
Valentin Tiede
Keyword(s):  
High Pressure ◽  
Flow Separation ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Centrifugal Impeller ◽  
Angle Distribution ◽  
Blade Angle ◽  
Minor Variation ◽  
Design Point ◽  
High Pressure Ratio

In an industrial research project of German and Swiss Turbo Compressor manufacturers a high pressure ratio centrifugal impeller was designed and investigated. Performance measurements and extensive laser measurements (L2F) of the flow field upstream, along the blade passage and downstream of the impeller have been carried out. In addition to that, 3D calculations have been performed, mainly for the design point. Results have been presented by Krain et al., 1995 and 1998, Eisenlohr et al., 1998 and Hah et al.,1999. During the design period of this impeller a radial blade at the inlet region was mandatory to avoid a rub at the shroud due to stress reasons. The measurements and the 3D calculations performed later, however, showed a flow separation at the hub near the leading edge due to too high incidence. Additionally a rather large exit width and a high shroud curvature near the exit caused a flow separation near the exit, which is enlarged by the radially transported wake of the already addressed hub separation. Changes to the hub blade angle distribution to reduce the hub incidence and an adaptation of the shroud blade angle distribution for the same impeller mass-flow at the design point were investigated by means of 3D calculations first with the same contours at hub and shroud; this was followed by calculations with a major change of the shroud contour including an exit width change with a minor variation of the hub contour. These calculations showed encouraging results; some of them will be presented in conjunction with the geometry data of the original impeller design.

The Matching of a Vaned Diffuser With a Radial Compressor Impeller and Its Effect on the Stage Performance

2014 ◽  
Vol 136 (12) ◽  
Author(s):  
Michael Casey ◽  
Daniel Rusch
Keyword(s):  
Test Data ◽  
Pressure Ratio ◽  
Prediction Method ◽  
Design Guidelines ◽  
Centrifugal Impeller ◽  
Vaned Diffuser ◽  
Radial Compressor ◽  
High Pressure Ratio ◽  
Wide Range ◽  
Matching Equation

The matching of a vaned diffuser with a centrifugal impeller is examined with a one-dimensional (1D) analysis combined with extensive experimental data. A matching equation is derived to define the required throat area of the diffuser relative to the throat area of the impeller at different design speeds and validated by comparison with a wide range of compressor designs. The matching equation is then used to give design guidelines for the throat area of vaned diffusers operating with impellers at different tip-speed Mach numbers. An analysis of test data for a range of high pressure ratio turbocharger compressor stages is presented in which different matching between the diffuser and the impeller has been experimentally examined. The test data includes different impellers with different diffuser throat areas over a wide range of speeds. It is shown that the changes in performance with speed and diffuser throat area can be explained on the basis of the tip-speed Mach number which causes both the diffuser and impeller to choke at the same mass flow. Based on this understanding, a radial compressor map prediction method is extended to include this parameter, so that more accurate maps for matched and mismatched vaned diffusers can be predicted.

The Matching of a Vaned Diffuser With a Radial Compressor Impeller and its Effect on the Stage Performance

2014 ◽  
Author(s):  
Michael Casey ◽  
Daniel Rusch
Keyword(s):  
Test Data ◽  
Pressure Ratio ◽  
Prediction Method ◽  
Design Guidelines ◽  
Centrifugal Impeller ◽  
Vaned Diffuser ◽  
Radial Compressor ◽  
High Pressure Ratio ◽  
Wide Range ◽  
Matching Equation

The matching of a vaned diffuser with a centrifugal impeller is examined with a one-dimensional (1D) analysis combined with extensive experimental data. A matching equation is derived to define the required throat area of the diffuser relative to the throat area of the impeller at different design speeds and validated by comparison with a wide range of compressor designs. The matching equation is then used to give design guidelines for the throat area of vaned diffusers operating with impellers at different tip-speed Mach numbers. An analysis of test data for a range of high pressure ratio turbocharger compressor stages is presented in which different matching between the diffuser and the impeller has been experimentally examined. The test data includes different impellers with different diffuser throat areas over a wide range of speeds. It is shown that the changes in performance with speed and diffuser throat area can be explained on the basis of the tip-speed Mach number which causes both the diffuser and impeller to choke at the same mass flow. Based on this understanding, a radial compressor map prediction method is extended to include this parameter, so that more accurate maps for matched and mismatched vaned diffusers can be predicted.

Aerodynamics and Aeroelasticity of Impeller Vane Interactions in a High Pressure Ratio Centrifugal Compressor

2010 ◽  
Author(s):  
S. K. Krishnababu ◽  
M. Imregun ◽  
J. S. Green ◽  
D. Hoyniak
Keyword(s):  
High Pressure ◽  
Centrifugal Compressor ◽  
Computational Study ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Computational Domain ◽  
Trailing Edge ◽  
Vaned Diffuser ◽  
High Pressure Ratio ◽  
Radial Gap

The interaction between impeller and diffuser in a high-pressure ratio centrifugal compressor is considered to have a strong influence on the unsteady flow field, the impeller response and the performance of the compressor. A computational study was performed to investigate the interactions between a backswept impeller and its downstream vaned diffuser with emphasis on the impeller response at 2 different vane settings. The unsteady computations were conducted using two different modelling levels of increasing fidelity. The computational domain included an impeller with 15 main and 15 splitter blades and 22-vane wedge diffuser. A steady-state stage calculation with a mixing-plane interface between the impeller trailing edge and the vane leading edge was conducted first to assess the performance. A whole-annulus unsteady stage calculation was conducted to study the response of the impeller. The effect of radial gap between the impeller trailing edge and the vane leading edge on the performance of the impeller was investigated in some detail. In agreement with other similar studies, the results suggest that there is an optimum value of the radius ratio for best performance.

scholarly journals Experimental Results on a Rotatable Low Solidity Vaned Diffuser

1992 ◽  
Author(s):  
J. M. Sorokes ◽  
J. P. Welch
Keyword(s):  
Pressure Ratio ◽  
Leading Edge ◽  
Centrifugal Impeller ◽  
Channel Inlet ◽  
Test Results ◽  
Test Vehicle ◽  
Vaned Diffuser ◽  
Critical Areas ◽  
Trailing Edges ◽  
Return Channel

This paper reviews test results from a rotatable low solidity vaned diffuser (RLSD). The device was installed in a single stage test vehicle consisting of a pseudo-return channel inlet, a subsonic centrifugal impeller (approx. 2:1 pressure ratio), the rotatable diffuser, a return channel and a dump collector. Static taps, total pressure probes, and thermocouples were located in critical areas throughout the stage. Manual traverse probes measured the pressure and angle profiles at RLSD leading and trailing edges. Results for various stagger angles, leading edge radius ratios, etc. are presented in terms of pressure recovery (Cp) and loss coefficient (LC). Comments are made regarding the applicability of the RLSD in production units.

Stall Behavior in an Ultra-High-Pressure-Ratio Centrifugal Compressor: Backward-Traveling Rotating Stall

2021 ◽  
pp. 1-51
Author(s):  
Yingjie Zhang ◽  
Xingen Lu ◽  
Yanfeng Zhang ◽  
Ziqing Zhang ◽  
Xu Dong ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Shock Intensity ◽  
Incidence Angle ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Boundary Layer Separation ◽  
Rotating Stall ◽  
High Pressure Ratio ◽  
Ultra High Pressure

Abstract This paper describes the stall mechanism in an ultra-high-pressure-ratio centrifugal compressor. A model comprising all impeller and diffuser blade passages is used to conduct unsteady simulations that trace the onset of instability in the compressor. Backward-traveling rotating stall waves appear at the inlet of the radial diffuser when the compressor is throttled. Six stall cells propagate circumferentially at approximately 0.7% of the impeller rotation speed. The detached shock of the radial diffuser leading edge and the number of stall cells determine the direction of stall propagation, which is opposite to the impeller rotation direction. Dynamic mode decomposition is applied to instantaneous flow fields to extract the flow structure related to the stall mode. This shows that intensive pressure fluctuations concentrate in the diffuser throat as a result of changes in the detached shock intensity. The diffuser passage stall and stall recovery are accompanied by changes in incidence angle and shock wave intensity. When the diffuser passage stalls, the shock-induced boundary-layer separation region near the diffuser vane suction surface gradually expands, increasing the incidence angle and decreasing the shock intensity. The shock is pushed from the diffuser throat toward the diffuser leading edge. When the diffuser passage recovers from stall, the shock wave gradually returns to the diffuser throat, with the incidence angle decreasing and the shock intensity increasing. Once the shock intensity reaches its maximum, the diffuser passage suffers severe shock-induced boundary-layer separation and stalls again.

Correlations Hidden in Compressor Maps

2011 ◽  
Author(s):  
Joachim Kurzke
Keyword(s):  
High Speed ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Operating Conditions ◽  
Specific Work ◽  
Gas Generator ◽  
High Pressure Ratio ◽  
Map Adaptation ◽  
Physical Phenomena ◽  
Corrected Flow

Realistic compressor maps are the key to high quality gas turbine performance calculations. When modeling the performance of an existing engine then these maps are usually not known and must be approximated by adapting maps from literature to either measured data or to other available information. There are many publications describing map adaptation processes, simple ones and more sophisticated physically based scaling rules. There are also reports about using statistics, genetic algorithms, neural networks and even morphing techniques for re-engineering compressor maps. This type of methods does not consider the laws of physics and consequently the generated maps are valid at best in the region in which they have been calibrated. This region is frequently very narrow, especially in case of gas generator compressors which run in steady state always on a single operating line. This paper describes which physical phenomena influence the shape of speed and efficiency lines in compressor maps. For machines operating at comparatively low speeds (so that the flow into each stage is subsonic), there is usually considerable range between choke and stall corrected flow. As the speed of the machine is increased the range narrows. For high-speed stages with supersonic relative flow into the rotor the efficiency maximum is where the speed line turns over from vertical to lower than maximum corrected flow. At this operating condition the shock is about to detach from the leading edge of the blades. The flow at a certain speed can also be limited by choking in the compressor exit guide vanes. For high pressure ratio single stage centrifugal compressors this is a normal case, but it can also happen with low pressure ratio multistage boosters of turbofan engines, for example. If the compressor chokes at the exit, then the specific work remains constant along the speed line while the overall pressure ratio varies and that generates a very specific shape of the efficiency contour lines in the map. Also in other parts of the map, the efficiency varies along speed lines in a systematic manner. Peculiar shapes of specific work and corrected torque lines can reveal physically impossibilities that are difficult to see in the standard compressor map pictures. Compressor maps generated without considering the inherent physical phenomena can easily result in misleading performance calculations if used at operating conditions outside of the region where they have been calibrated. Whatever map adaptation method is used: the maps created in such a way should be checked thoroughly for violations of the underlying laws of compressor physics.

Detailed Flow Study of Mach Number 1.6 High Transonic Flow in a Pressure Ratio 11 Centrifugal Compressor Impeller: Part 2 — Effect of the Inducer Shroud Bleed and Development of a Low Energy Region Along the Shroud

2007 ◽  
Author(s):  
Hirotaka Higashimori ◽  
Susumu Morishita ◽  
Masayuki Suzuki ◽  
Tooru Suita
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Centrifugal Compressor ◽  
Transonic Flow ◽  
Energy Region ◽  
Pressure Ratio ◽  
Low Energy ◽  
Actual Measurement ◽  
High Pressure Ratio ◽  
Compressor Impeller

Requirements for aeronautical gas turbine engines for helicopters include small size, low weight, high output, and low fuel consumption. In order to achieve these requirements, development work has been carried out on high pressure ratio compressors with high efficiency. As a result, we have developed a single stage centrifugal compressor with a pressure ratio of 11 for a 1000 shp class gas turbine. This report presents a study on the internal flow of a high pressure ratio centrifugal compressor impeller. The centrifugal compressor is a high transonic compressor with an inlet Mach number of about 1.6. In high inlet Mach number compressors, the flow in the inducer is a complex transonic flow characterized by interaction between the shockwave and boundary layer, while the flow in the middle of the impeller is a distorted flow with a low energy region. In order to ensure the reliability of aerodynamic design technology for such transonic centrifugal compressors, the complex transonic flow and formation of the low energy region predicted by CFD must be actually measured, comparison must be undertaken between the CFD results and the actual flow measurement, and the accuracy and other issues pertaining to CFD must be clarified. In a previous report [12], we elucidated the flow in the inducer of a high transonic impeller by means of LDV and unsteady pressure measurement. That report showed that, in the flow of an inducer with a Mach number of approx. 1.6, the oblique shockwave in the middle of the impeller throat interacts with the blade tip leakage flow, and that reverse flow occurs in the vicinity of the casing. Furthermore, although CFD predicted a low energy region in the splitter portion, this could not be detected in actual measurement. In the context of the current report, comparative verification of the CFD and LDV measurement results was undertaken with respect to the formation of the casing wall surface boundary layer in the transonic flow within the inducer. In this conjunction, inducer bleed was introduced to control this boundary layer, and the effect of the inducer bleed on the flow was ascertained through actual measurement. It was also sought to additionally confirm the “low energy region” in the splitter. Accordingly, the flow velocity distribution was measured at two sections, thereby clarifying the characteristics of the actual flow in the region. The impeller for which measurement was performed has the same specifications as that in the previous report (see Table 1). In the present report, so as to measure the flow under conditions encouraging the formation of a boundary layer accompanying substantial inducer deceleration, measurement was conducted at 95% of design speed and a relative Mach number at the blade tips of about 1.5.

Aerodynamic Analysis and Multi-Objective Optimization Design of a High Pressure Ratio Centrifugal Impeller

2014 ◽  
Author(s):  
Zhendong Guo ◽  
Zhiming Zhou ◽  
Liming Song ◽  
Jun Li ◽  
Zhenping Feng
Keyword(s):  
Data Mining ◽  
High Pressure ◽  
Pressure Ratio ◽  
Differential Evolution Algorithm ◽  
Aerodynamic Performance ◽  
Centrifugal Impeller ◽  
Tip Leakage Flow ◽  
Pareto Solutions ◽  
Multi Objective ◽  
High Pressure Ratio

The design of high pressure ratio impellers is a challenging task. SRV2-O, a typical high pressure ratio centrifugal impeller is selected for the research. A good understanding of flow characteristics in the passage of SRV2-O is obtained by using 3D Reynolds-Averaged Navier-Stokes (RANS) solutions upon numerical validation. It confirms that tip leakage flow and shock wave boundary layer interactions produce the primary energy loss in this transonic impeller. A 3D multi-objective aerodynamic optimization and data mining method named BMOE is presented and programmed by integrating a self-adaptive multi-objective differential evolution algorithm SMODE, 3D blade parameterization method based on non-uniformed B-Spline, RANS solver technique and self-organization map (SOM) based data mining technique. Using BMOE, multi-objective aerodynamic design optimization and data mining is performed for SRV2-O. 14 Pareto solutions are obtained for maximizing isentropic efficiency and total pressure ratio of the impeller. Three typical Pareto solutions, Design A with the highest efficiency, Design B with the higher efficiency and larger pressure ratio and Design C with the maximum pressure ratio, are analyzed. Detailed analysis indicates that the aerodynamic performance of optimized designs is greatly improved. Furthermore, by SOM-based data mining on optimization results, trade-off relation between objective functions and parameter influence mechanism on impeller aerodynamic performance are visualized and explored.

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