High Specific Speed, High Inducer Tip Mach Number, Centrifugal Compressor

2003 ◽  
Author(s):  
C. Rodgers
Keyword(s):  
Mach Number ◽  
Centrifugal Compressor ◽  
Transonic Flow ◽  
High Efficiency ◽  
Flow Fields ◽  
Test Results ◽  
Positive Slope ◽  
Stall Margin ◽  
Recovery Characteristic ◽  
Specific Speed

The demand for more efficient turbocharger and aviation centrifugal compressors operating at higher pressure ratios and specific speeds with extended flow ranges is focusing research efforts on the inducer and diffuser transonic flow fields. At pressure ratios above 5.0 and specific speeds of unity inducer tip relative Mach numbers exceeding 1.4 can be encountered, precipitating both increased shock losses and diminished stall margin. The results of compressor rig testing on a research 6.8 inch tip (173mm) diameter single stage centrifugal compressor operating with inducer tip relative Mach number up to 1.5 are presented. The test results reveal high efficiency combined with extended flow range. This was achieved through improved impeller stability with shroud bleed, thereby permitting the diffuser to operate stably on its positive slope recovery characteristic.

Performance Analysis of the Test Results on a Two-Stage Transonic Fan

1983 ◽  
Vol 105 (1) ◽  
pp. 125-129
Author(s):  
Baoshi Chen ◽  
Tianyi Zhang
Keyword(s):  
Mach Number ◽  
Performance Analysis ◽  
Radial Distribution ◽  
Flow Separation ◽  
High Load ◽  
Test Results ◽  
Two Stage ◽  
Stall Margin ◽  
High Incidence ◽  
Transonic Fan

Test results obtained from a two-stage fan are analysed and the reasons that caused the design performance target not to be attained are presented in this paper. Addition of a partspan shroud on rotor 1 caused higher losses and changed radial distribution of parameters. Modification on the flowpath and chord length of stator 1 resulted in excessively high inlet Mach number and flow separation in the hub region. The high load and high incidence at the hub of rotor 2 caused higher losses and reduced stall margin of the fan.

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.

Detailed Flow Study of Mach Number 1.6 High Transonic Flow With a Shock Wave in a Pressure Ratio 11 Centrifugal Compressor Impeller

2004 ◽  
Vol 126 (4) ◽  
pp. 473-481 ◽  
Author(s):  
Hirotaka Higashimori ◽  
Kiyoshi Hasagawa ◽  
Kunio Sumida ◽  
Tooru Suita
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Mach Number ◽  
Centrifugal Compressor ◽  
Transonic Flow ◽  
Reverse Flow ◽  
Pressure Ratio ◽  
Layer Interaction ◽  
Shock Boundary Layer Interaction ◽  
Boundary Layer Interaction

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 efficiency and high pressure ratio compressors. As a result, we have developed a single stage centrifugal compressor with a pressure ratio of 11 for a 1000 shp class gas turbine. 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 distortion due to the shock wave and the shock boundary layer interaction must have a large effect on the flow in the inducer. In order to ensure the reliability of aerodynamic design technology, the actual supersonic flow phenomena with a shock wave must be ascertained using measurement and Computational Fluid Dynamics (CFD). This report presents the measured results of the high transonic flow at the impeller inlet using Laser Doppler Velocimeter (LDV) and verification of CFD, with respect to the high transonic flow velocity distribution, pressure distribution, and shock boundary layer interaction at the inducer. The impeller inlet tangential velocity is about 460 m/s and the relative Mach number reaches about 1.6. Using a LDV, about 500 m/s relative velocity was measured preceding a steep deceleration of velocity. The following steep deceleration of velocity at the middle of blade pitch clarified the cause as being the pressure rise of a shock wave, through comparison with CFD as well as comparison with the pressure distribution measured using a high frequency pressure transducer. Furthermore, a reverse flow is measured in the vicinity of casing surface. It was clarified by comparison with CFD that the reverse flow is caused by the shock-boundary layer interaction. Generally CFD shows good agreement with the measured velocity distribution at the inducer and splitter inlet, except in the vicinity of the casing surface.

Detailed Flow Study of Mach Number 1.6 High Transonic Flow With a Shock Wave in a Pressure Ratio 11 Centrifugal Compressor Impeller

2004 ◽  
Author(s):  
Hirotaka Higashimori ◽  
Kiyoshi Hasagawa ◽  
Kunio Sumida ◽  
Tooru Suita
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Mach Number ◽  
Centrifugal Compressor ◽  
Transonic Flow ◽  
Reverse Flow ◽  
Pressure Ratio ◽  
Layer Interaction ◽  
Shock Boundary Layer Interaction ◽  
Boundary Layer Interaction

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 efficiency and high pressure ratio compressors. As a result, we have developed a single stage centrifugal compressor with a pressure ratio of 11 for a 1000 shp class gas turbine. 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 distortion due to the shock wave and the shock boundary layer interaction must have a large effect on the flow in the inducer. In order to ensure the reliability of aerodynamic design technology, the actual supersonic flow phenomena with a shock wave must be ascertained using measurement and CFD. This report presents the measured results of the high transonic flow at the impeller inlet using LDV and verification of CFD, with respect to the high transonic flow velocity distribution, pressure distribution and shock boundary layer interaction at the inducer. The impeller inlet tangential velocity is about 460m/s and the relative Mach number reaches about 1.6. Using an LDV, about 500m/s relative velocity was measured preceding a steep deceleration of velocity. The following steep deceleration of velocity at the middle of blade pitch clarified the cause as being the pressure rise of a shock wave, through comparison with CFD as well as comparison with the pressure distribution measured using a high frequency pressure transducer. Furthermore, a reverse flow is measured in the vicinity of casing surface. It was clarified by comparison with CFD that the reverse flow is caused by the shock-boundary layer interaction. Generally CFD shows good agreement with the measured velocity distribution at the inducer and splitter inlet, except in the vicinity of the casing surface.

A Diffusion Factor Correlation for Centrifugal Impeller Stalling

1978 ◽  
Vol 100 (4) ◽  
pp. 592-601 ◽  
Author(s):  
C. Rodgers
Keyword(s):  
Centrifugal Compressor ◽  
High Loading ◽  
Centrifugal Impeller ◽  
Preliminary Design ◽  
Test Results ◽  
Diffusion Factor ◽  
Speed Range ◽  
Parallel Wall ◽  
Specific Speed ◽  
Vaneless Diffusers

Test results pertaining to the stalling characteristics of centrifugal compressor impellers with parallel wall vaneless diffusers are presented and studied to correlate the coincidence of stall with a limiting impeller diffusion capability. It is suggested that a modified diffusion factor, to include the effects of meridional curvature, provides improved stall correlation for a wide specific speed range of backswept impeller types. The possibility of applying this diffusion factor to high loading radially bladed impellers is discussed as dependent upon blockage and windage plus recirculation effects. Use of the diffusion factor limit in the preliminary design of most common turbomachinery types, incompressible and compressible, to assess impeller (or rotor) stall is conceivable.

Using CFD to Improve Stall Margin

1999 ◽  
Author(s):  
James R. Hardin ◽  
Charles F. Boal
Keyword(s):  
High Efficiency ◽  
Industrial Applications ◽  
Stable Operation ◽  
Test Results ◽  
Peak Efficiency ◽  
Centrifugal Compressors ◽  
Stall Margin ◽  
Operating Range ◽  
Multi Stage ◽  
Cfd Analyses

Abstract Centrifugal compressors for multi-stage industrial applications must have both high efficiency and stable operation over a wide flow range. CFD analyses were used to evaluate the stall margins of candidate impeller designs compared to a baseline semi-inducer design. With this guidance, a new full-inducer impeller was designed with predicted wider operating range and slightly higher peak efficiency. Prototype tests confirmed these predictions. This paper presents predictions from two CFD codes and test results, for both the original and the redesigned impeller, and discusses the application of CFD for predicting impeller stall.

Three-Dimensional Rotational Flow in Transonic Turbomachines: Part II — Full 3D Flow in CAS Rotor Obtained by Using a Number of S1 and S2 Stream Filaments

1990 ◽  
Author(s):  
Wu Chung-Hua ◽  
Zhao Xiaolu ◽  
Qin Lisen
Keyword(s):  
Mach Number ◽  
Transonic Flow ◽  
Three Dimensional ◽  
Flow Fields ◽  
Rotational Flow ◽  
Three Dimensional Flow ◽  
Measured Velocity ◽  
Dimensional Flow ◽  
Iterative Calculation ◽  
Good Agreement

The general theory for three–dimensional flow in subsonic and supersonic turbomachines has recently been extended to transonic turbomachines. In Part II of the paper, quasi– and full three–dimensional solutions of the transonic flow in the CAS rotor are presented. The solutions are obtained by iterative calculation between a number of S1 stream filaments and, respectively, a central S2 stream filament and a number of S2m stream filaments. Relatively simple methods developed recently for solving the transonic flow along S1 and S2 stream filaments are used in the calculation. The three–dimensional flow fields in the CAS rotor obtained by the present method are presented in detail with special emphasis on the converging process for the configuration of the S1 and S2 stream filaments. The three–dimensional flow fields obtained in the quasi– and full 3D solutions are quite similar, but the former gives a lower peak Mach number and a smaller circumferential variation in Mach number than the latter. A comparison between the theoretical solution and the Laser–2–Focus measurement shows that the character of the transonic flow including the 3D shock structure is in good agreement, but the measured velocity is slightly higher than the calculated one over most of the flow field.

Design of a Centrifugal Compressor With Low Specific Speed for Automotive Fuel Cell

2008 ◽  
Author(s):  
Xinqian Zheng ◽  
Yangjun Zhang ◽  
Hong He ◽  
Zhiling Qiu
Keyword(s):  
Fuel Cell ◽  
Centrifugal Compressor ◽  
Electric Motor ◽  
High Speed ◽  
High Efficiency ◽  
Pressure Ratio ◽  
System A ◽  
Automotive Fuel ◽  
Low Specific Speed ◽  
Specific Speed

Centrifugal compressors driven by electric motor are the promising type for fuel cell pressurization system. A low specific speed centrifugal compressor powered by an ordinary high-speed (about 25,000rpm) electric motor has been designed at Tsinghua University for automotive fuel cell engines. The experimental results indicate that the designed low specific speed centrifugal compressor has comparatively high efficiency and wide operating range. In the condition of designed speed (24,000rpm), the highest efficiency and pressure ratio of the centrifugal compressor is up to 70% and 1.6, respectively. The designed low specific speed centrifugal compressor can meet the requirement of air systems of automotive fuel cell engines preliminarily. Moreover, the low specific speed centrifugal compressor avoids difficulties of usage of ultra-high-speed electric motors (about 60,000rpm) in high specific speed compressor. Based on the preliminary results of this centrifugal compressor, a new low specific speed centrifugal compressor with higher performances is being developed.

Development of Loss Correlation and Tool Validation at Transonic Condition Based on Cascade Test

2021 ◽  
Author(s):  
Jaewoo Choi ◽  
David Simurda ◽  
Jaewook Song ◽  
Martin Luxa ◽  
Sungryong Lee ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Transonic Flow ◽  
High Efficiency ◽  
Thickness Distribution ◽  
Leading Edge ◽  
Suction Side ◽  
Loss Coefficient ◽  
Operating Range ◽  
Distribution Type

Abstract Overall efficiency of an axial compressor is largely affected by its front stage when it is operating under transonic flow conditions. For this reason, many manufacturers and researchers are advancing research and development of transonic airfoils in these days. Doosan, in frame of a development of high efficiency gas turbine, developed high efficiency airfoil for a transonic rotor and conducted cascade tests. Therefore, this study deals with a test of two compressor transonic blade cascades at inlet Mach number over 1.1. To improve the efficiency and operating range, two kinds of thickness distribution type based on Enhanced Doosan Airfoil (EDA), which applied unique rule, were applied and assessed. The first airfoil consists of polynomial thickness distribution and the second airfoil consists of new thickness distribution with specially tailored leading edge. In order to ensure accurate geometry of a model, a detailed checkout process upon production of model blades used in the test was performed. This is because, in the case of transonic airfoil, if the inlet leading edge shape differs by more than 0.2% than designed airfoil of leading edge, the result will be completely different. Therefore, not only the tolerance within 0.1% was confirmed but also the shape produced through simulation and 3D CMM scan data. The main parameters for the comparison are an inlet Mach number, an axial velocity density ratio (AVDR) and the kind of thickness distribution. Results of tests and CFD blade to blade analysis using MISES 2.70 are compared. The flow field was visualized using schlieren technique and parameters of the suction side boundary layer were evaluated at several locations based on Pitot probe traverses. The results confirm that a suction peak at the round leading edge disappears in the case of the new thickness type distribution with tailored leading edge. This confirms that the profile shaping without jump in curvature in the leading edge region leads to smooth acceleration without peaks. Nevertheless, results show that the new thickness distribution type is not absolutely good in comparison with the polynomial thickness distribution type with respect to the total pressure loss coefficient. Moreover, bucket range (operating range) is also almost the same. Results of the suction side boundary layer traversing suggest that the transition of the boundary layer takes place beyond the location x/cax > 0.088. The MISES results show that a shock location and the boundary layer parameters are similar to test results. However, values of the loss coefficient show some difference. Therefore, a new correlation in particular transonic flow condition was developed.

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