Design Consideration of a Centrifugal Compressor Impeller With High Mach Number and High Swirl Angle at Exit

2010 ◽  
Author(s):  
Kangsoo Im
Keyword(s):  
Mach Number ◽  
Centrifugal Compressor ◽  
Swirl Flow ◽  
Stable Operation ◽  
Compression Process ◽  
Vaned Diffuser ◽  
Flow Angle ◽  
Compressor Impeller ◽  
High Mach ◽  
Exit Mach Number

A centrifugal compressor impeller with high exit Mach number (M = 0.8–1.1) & high outlet flow angle (ALPHA 2 = 78–85 degrees) was designed. The objective of this work is to create a real condition for the better understanding of the behavior of the fluid flow in transonic and high swirl flow vaned diffusers. Particular attention was given to the flows located at the rotor–stator interface. The unsteady interaction between impeller and diffuser plays an important role in the compression process, especially in high loaded impellers. Actually, the vaned diffuser has to tolerate the distorted upstream flow due to the jet-wake structure coming from the impeller. Following the impeller design, numerical investigations were conducted for the transonic centrifugal compressor impeller composed of backswept unshrouded blades. The characteristic plots of the impeller resulting from the simulations show a good stable operation.

High Inlet Relative Mach Number Centrifugal Compressor Impeller Design

2007 ◽  
Author(s):  
James M. Sorokes ◽  
Jason A. Kopko
Keyword(s):  
Mach Number ◽  
Test Data ◽  
Centrifugal Compressor ◽  
High Mach Number ◽  
Performance Map ◽  
Compressor Impeller ◽  
The Subject ◽  
Mach Numbers ◽  
High Mach

This document presents an overview of impeller inlet relative Mach number, how the parameter is calculated, and its importance as an indicator of impeller performance. Comments are also offered regarding the comparison of inlet relative Mach numbers obtained from different compressor vendors. A sample impeller is used to illustrate the various methods used to calculate the inlet relative Mach number. Test data for that impeller is also offered to indicate the performance map achievable with high Mach number designs. Please note that this document is not intended to be an all-inclusive treatment of the subject; rather, it summarizes the OEM’s methodologies and perspective.

Some Studies on Low Solidity Vaned Diffusers of a Centrifugal Compressor Stage

2005 ◽  
Author(s):  
T. Ch. Siva Reddy ◽  
G. V. Ramana Murty ◽  
Prasad Mukkavilli ◽  
D. N. Reddy
Keyword(s):  
Centrifugal Compressor ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Pressure Recovery ◽  
Compressor Stage ◽  
Recovery Coefficient ◽  
Vaned Diffuser ◽  
Flow Angle ◽  
Centrifugal Compressor Stage ◽  
Pressure Recovery Coefficient

Numerical simulation of impeller and low solidity vaned diffuser (LSD) of a centrifugal compressor stage is performed individually using CFX- BladeGen and BladeGenPlus codes. The tip mach number for the chosen study was 0.35. The same configuration was used for experimental investigation for a comparative study. The LSD vane is formed using standard NACA profile with marginal modification at trailing edge. The performance parameters obtained form numerical studies at the exit of impeller and the diffuser have been compared with the corresponding experimental data. These parameters are pressure ratio, polytropic efficiency and flow angle at the impeller exit where as the parameters those have been compared at the exit of diffuser are the static pressure recovery coefficient and the exit flow angle. In addition, the numerical prediction of the blade loading in terms of blade surface pressure distribution on LSD vane has been compared with the corresponding experimental results. Static pressure recovery coefficient and flow angle at diffuser exit is seen to match closely at higher flows. The difference at lower flows could be due to the effect of interaction between impeller and diffuser combinations, as the numerical analysis was done separately for impeller and diffuser and the effect of impeller diffuser interaction was not considered.

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.

Effects of Inlet Flow Field Conditions on the Performance of Centrifugal Compressor Diffusers: Part 2 — Straight-Channel Diffuser

1998 ◽  
Author(s):  
Sabri Deniz ◽  
Edward M. Greitzer ◽  
Nicholas A. Cumpsty
Keyword(s):  
Mach Number ◽  
Flow Field ◽  
Centrifugal Compressor ◽  
Axial Flow ◽  
Rotating Stall ◽  
Pressure Recovery ◽  
Straight Channel ◽  
Inlet Flow ◽  
Flow Angle ◽  
Operating Range

This is Part 2 of an examination of influence of inlet flow conditions on the performance and operating range of centrifugal compressor vaned diffusers. The paper describes tests of straight-channel type diffuser, sometimes called a wedge-vane diffuser, and compares the results with those from the discrete-passage diffusers described in Part 1. Effects of diffuser inlet Mach number, flow angle, blockage, and axial flow non-uniformity on diffuser pressure recovery and operating range are addressed. The straight-channel diffuser investigated has 30 vanes and was designed for the same aerodynamic duty as the discrete-passage diffuser described in Part 1. The ranges of the overall pressure recovery coefficients were 0.65–0.78 for the straight-channel diffuser and 0.60–0.70 for the discrete-passage diffuser; the pressure recovery of the straight-channel diffuser was roughly 10% higher than that of the discrete-passage diffuser. Both types of the diffusers showed similar behavior regarding the dependence on diffuser inlet flow angle and the insensitivity of the performance to inlet flow field axial distortion and Mach number. The operating range of the straight-channel diffuser, as for the discrete-passage diffusers was limited by the onset of rotating stall at a fixed momentum-averaged flow angle into the diffuser, which was for the straight-channel diffuser, αcrit = 70° ±0.5°. The background, nomenclature and description of the facility and method are all given in Part 1.

scholarly journals Performance Improvement of Return Channel in Multistage Centrifugal Compressor Using Multi-Objective Optimization

2012 ◽  
Author(s):  
Y. Nishida ◽  
H. Kobayashi ◽  
H. Nishida ◽  
K. Sugimura
Keyword(s):  
Centrifugal Compressor ◽  
Pressure Coefficient ◽  
Swirl Flow ◽  
Optimized Design ◽  
Original Design ◽  
Flow Angle ◽  
Second Stage ◽  
Return Channel ◽  
Multistage Centrifugal Compressor ◽  
Stage Efficiency

The effect of the design parameters of a return channel on the performance of a multistage centrifugal compressor was numerically investigated, and the shape of the return channel was optimized using a multi-objective optimization method based on a genetic algorithm to improve the performance of the centrifugal compressor. The results of sensitivity analysis using Latin hypercube sampling suggested that the inlet-to-outlet area ratio of the return vane affected the pressure loss in the return channel, and that the inlet-to-outlet radius ratio of the return vane affected the outlet flow angle from the return vane. Moreover, this analysis suggested that the number of return vanes affected both the loss and the flow angle at the outlet. As a result of optimization, the number of return vanes was increased from 14 to 22 and the area ratio was decreased from 0.71 to 0.66. The radius ratio was also decreased from 2.1 to 2.0. Performance tests on a centrifugal compressor with two return channels (the original design and optimized design) were carried out using two-stage test apparatus. The measured flow distribution exhibited a swirl flow in the center region and a reversed swirl flow near the hub and shroud sides. The exit flow of the optimized design was more uniform than that of the original design. For the optimized design, the overall two-stage efficiency and pressure coefficient were increased by 0.7% and 1.5%, respectively. Moreover, the second-stage efficiency and pressure coefficient were respectively increased by 1.0% and 3.2%, It is considered that the increase in the second-stage efficiency was caused by the increased uniformity of the flow, and the rise in the pressure coefficient was caused by a decrease in the residual swirl flow. It was thus concluded from the numerical and experimental results that the optimized return channel improved the performance of the multistage centrifugal compressor.

Impeller Rotating Stall as a Trigger for the Transition From Mild to Deep Surge in a Subsonic Centrifugal Compressor

1993 ◽  
Author(s):  
Beat Ribi ◽  
Georg Gyarmathy
Keyword(s):  
Mach Number ◽  
Mass Flow Rate ◽  
Flow Rate ◽  
Centrifugal Compressor ◽  
Critical Value ◽  
Rotating Stall ◽  
Vaned Diffuser ◽  
Time Resolved ◽  
Compressor Map ◽  
Operating Points

The present paper concerns the transition from mild to deep surge in a single stage centrifugal compressor using a vaned diffuser. Time-resolved measurements of the mass flow rate and the static pressures at various locations of the compressor were analyzed for different diffuser geometries and different operating points in the compressor map. When the throttle valve was gradually closed at an impeller tip Mach number (Mu) above 0.4, the compressor showed first mild and then deep surge whereas at Mu=0.4 rotating stall was the dominant instability. This single-cell rotating stall was identified to be caused by the impeller. During mild surge at higher Mach numbers the instantaneous flow and pressure traces showed that the overall flow at the stage inlet intermittently dropped below the critical value associated with the occurence of impeller rotating stall. Rotating stall appeared for a while but vanished as soon as the flow increased again. With further throttling, however, a threshold was reached at which rotating stall triggered deep surge. The results show that triggering only occurred if the flow deficiency causing rotating stall persisted long enough to permit the stall cell to make at least one or two revolutions.

Investigation of an Inversely Designed Centrifugal Compressor Stage: Part 1 — Design and Numerical Verification

2003 ◽  
Author(s):  
M. Zangeneh ◽  
M. Schleer ◽  
F. Plo̸ger ◽  
S. S. Hong ◽  
C. Roduner ◽  
...  
Keyword(s):  
Centrifugal Compressor ◽  
Design Method ◽  
Leading Edge ◽  
Inverse Design ◽  
Numerical Verification ◽  
Appreciable Effect ◽  
Tip Leakage Flow ◽  
Flow Angle ◽  
Compressor Impeller ◽  
Blade Geometry

In this paper the 3D inverse design code TURBOdesign-1 is applied to the design of the blade geometry of a centrifugal compressor impeller with splitter blades. In the design of conventional impellers the splitter blades normally have the same geometry as the full blades and are placed at mid-pitch location between the two full blades, which can usually result in a mis-match between the flow angle and blade angles at the splitter leading edge. In the inverse design method the splitter and full blade geometry is computed independently for a specified distribution of blade loading on the splitter and full blades. In this paper the basic design methodology is outlined and then the flow in the conventional and inverse designed impeller is compared in detail by using CFD code TASCflow. The CFD results confirm that the inverse design impeller has a more uniform exit flow, better control of tip leakage flow and higher efficiency than the conventional impeller. The results also show that the shape of the trailing edge geometry has a very appreciable effect on the impeller Euler head and this must be accurately modeled in all CFD computations to ensure closer match between CFD and experimental results. Detailed measurements are presented in part 2 of the paper.

Off-Design Flow Measurements in a Centrifugal Compressor Vaneless Diffuser

1995 ◽  
Vol 117 (4) ◽  
pp. 602-608 ◽  
Author(s):  
A. Pinarbasi ◽  
M. W. Johnson
Keyword(s):  
Flow Rate ◽  
Centrifugal Compressor ◽  
Secondary Flows ◽  
Design Flow ◽  
Vaneless Diffuser ◽  
Flow Measurements ◽  
Vaned Diffuser ◽  
Flow Angle ◽  
Flow Rates ◽  
Study Results

Detailed measurements have been taken of the three-dimensional velocity field within the vaneless diffuser of a backswept low speed centrifugal compressor using hot-wire anemometry. A 16 percent below and an 11 percent above design flow rate were used in the present study. Results at both flow rates show how the blade wake mixes out more rapidly than the passage wake. Strong secondary flows inherited from the impeller at the higher flow rate delay the mixing out of the circumferential velocity variations, but at both flow rates these circumferential variations are negligible at the last measurement station. The measured tangential/radial flow angle is used to recommend optimum values for the vaneless space and vane angle for design of a vaned diffuser.

Flow Instability at the Inlet of a Centrifugal Compressor

1981 ◽  
Vol 103 (2) ◽  
pp. 451-456
Author(s):  
G. Flueckiger ◽  
A. Melling
Keyword(s):  
Centrifugal Compressor ◽  
Flow Instability ◽  
Laser Doppler Anemometry ◽  
Suction Side ◽  
Operating Conditions ◽  
Gas Velocity ◽  
Unstable Flow ◽  
Flow Angle ◽  
Compressor Impeller ◽  
Service Conditions

Using laser Doppler anemometry, two components of the gas velocity have been measured at the inlet of a centrifugal compressor impeller, operated at speeds typical of service conditions for a medium-sized turbocharger. The flow was found to be unstable, especially adjacent to the suction side of the blades, such that two predominant conditions existed in the flow. The unstable flow is illustrated in the paper by distributions of relative velocity and relative flow angle, and the effects of different operating conditions on these distributions are examined. The instability is believed to be caused by a pre-stall condition as the compressor operating point approaches a fully stalled condition which occurs during surge.

scholarly journals Investigation of the effects of main geometric parameters and flow characteristics on secondary flow losses in a turbine cascade

2021 ◽  
Vol 2131 (3) ◽  
pp. 032081
Author(s):  
M Mesbah ◽  
V G Gribin ◽  
K Souri
Keyword(s):  
Mach Number ◽  
Aspect Ratio ◽  
Total Energy ◽  
Flow Characteristics ◽  
Numerical Results ◽  
Energy Losses ◽  
Inlet Flow ◽  
Flow Angle ◽  
Flow Losses ◽  
Exit Mach Number

Abstract This paper presents numerical simulation results of a three-dimensional (3D) transitional flow in a stator cascade of an axial turbine. The influences of the main geometric parameters and flow characteristics including, the blade aspect ratio, pitch-to-chord ratio, inlet flow angle, and exit Mach number, on secondary flows development and end-wall losses, were studied. The numerical results were validated by the results of experiments conducted in the laboratory of the steam and gas turbine faculty of the Moscow Power Engineering Institute. The maximum difference between computed and experimental results was 2.4 %. The total energy losses decrease by 20 % when the exit Mach number changes from 0.38 to 0.8. Numerical results indicated that the blade aspect ratio had the most effect on secondary flow losses. The total energy losses increase by 46.6 % when the aspect ratio decreases from 1 to 0.25. The total loss of energy by 13.2 % decreases by increasing the inlet flow angle from 60 degrees to 90 degrees. Then by increasing the inlet flow angle from 90 to 110 degrees, the total loss rises by 3.6%. As the pitch-to-chord ratio increases from 0.7 to 0.75, the total energy losses are reduced by 12.2 %. Then by increasing the pitch-to-chord ratio from 0.75 to 0.8, the total energy losses increase by 6 %. As with experimental data, the numerical results showed that the optimal inlet flow angle and relative pitch for the cascade are 90 degrees and 0.75, respectively.

Sign in / Sign up

Export Citation Format

Share Document