Performance of Three Vaned Radial Diffusers With Swirling Transonic Flow

1975 ◽  
Vol 97 (2) ◽  
pp. 155-160 ◽  
Author(s):  
S. Baghdadi ◽  
A. T. McDonald
Keyword(s):  
Transonic Flow ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Motion Pictures ◽  
Back Pressure ◽  
Pressure Measurements ◽  
Performance Measurements ◽  
High Pressure Ratio ◽  
Compressor Impeller ◽  
And Performance

A unique vortex nozzle facility has been conceived and developed to simulate the exit flow from a high pressure ratio centrifugal compressor impeller. Visual studies and performance measurements have been made for three vane sets representing common designs for vaned radial diffusers. Motion pictures show the progression from choke through operating to surge conditions as the back pressure on the diffuser is increased. The films, together with total and static pressure measurements, indicate that surge is an instability triggered by flow separation in the vaneless or quasivaneless space ahead of the diffuser throat. A geometrical criterion for the onset of surge is identified. The surge-to-choke operating range of the three diffusers appears to be a function of the number of diffuser vanes only.

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.

scholarly journals The Design and Evaluation of a High Pressure Ratio Radial Turbine

1992 ◽  
Author(s):  
K. R. Pullen ◽  
N. C. Baines ◽  
S. H. Hill
Keyword(s):  
High Pressure ◽  
High Speed ◽  
Pressure Ratio ◽  
Performance Measurements ◽  
Turbine Engine ◽  
High Pressure Ratio ◽  
Turbine Efficiency ◽  
Wide Range ◽  
Dimensional Parameters ◽  
Very High

A single stage, high speed, high pressure ratio radial inflow turbine was designed for a single shaft gas turbine engine in the 200 kW power range. A model turbine has been tested in a cold rig facility with correct simulation of the important non-dimensional parameters. Performance measurements over a wide range of operation were made, together with extensive volute and exhaust traverses, so that gas velocities and incidence and deviation angles could be deduced. The turbine efficiency was lower than expected at all but the lowest speed. The rotor incidence and exit swirl angles, as obtained from the rig test data, were very similar to the design assumptions. However, evidence was found of a region of separation in the nozzle vane passages, presumably caused by a very high curvature in the endwall just upstream of the vane leading edges. The effects of such a separation are shown to be consistent with the observed performance.

scholarly journals Interpreting Aerodynamics of a Transonic Impeller from Static Pressure Measurements

2018 ◽  
Vol 2018 ◽  
pp. 1-9
Author(s):  
Fangyuan Lou ◽  
John Charles Fabian ◽  
Nicole Leanne Key
Keyword(s):  
Flow Field ◽  
High Speed ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Pressure Measurements ◽  
High Loss ◽  
Supersonic Inlet ◽  
Mach Numbers ◽  
Inlet Conditions

This paper investigates the aerodynamics of a transonic impeller using static pressure measurements. The impeller is a high-speed, high-pressure-ratio wheel used in small gas turbine engines. The experiment was conducted on the single stage centrifugal compressor facility in the compressor research laboratory at Purdue University. Data were acquired from choke to near-surge at four different corrected speeds (Nc) from 80% to 100% design speed, which covers both subsonic and supersonic inlet conditions. Details of the impeller flow field are discussed using data acquired from both steady and time-resolved static pressure measurements along the impeller shroud. The flow field is compared at different loading conditions, from subsonic to supersonic inlet conditions. The impeller performance was strongly dependent on the inducer, where the majority of relative diffusion occurs. The inducer diffuses flow more efficiently for inlet tip relative Mach numbers close to unity, and the performance diminishes at other Mach numbers. Shock waves emerging upstream of the impeller leading edge were observed from 90% to 100% corrected speed, and they move towards the impeller trailing edge as the inlet tip relative Mach number increases. There is no shock wave present in the inducer at 80% corrected speed. However, a high-loss region near the inducer throat was observed at 80% corrected speed resulting in a lower impeller efficiency at subsonic inlet conditions.

Investigation of Self-Recycling-Casing-Treatment (SRCT) Influence on Stability of High Pressure Ratio Centrifugal Compressor With a Volute

2011 ◽  
Author(s):  
Mingyang Yang ◽  
Ricardo Martinez-Botas ◽  
Yangjun Zhang ◽  
Xinqian Zheng ◽  
Takahiro Bamba ◽  
...  
Keyword(s):  
High Pressure ◽  
Numerical Method ◽  
Flow Rate ◽  
Centrifugal Compressor ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Flow Distortion ◽  
Casing Treatment ◽  
High Pressure Ratio ◽  
The Stability

Large feasible operation range is a challenge for high pressure ratio centrifugal compressor of turbocharger in vehicle engine. Self-Recycling-Casing-Treatment (SRCT) is a widely used flow control method to enlarge the range for this kind of compressor. This paper investigates the influence of symmetrical/asymmetrical SRCT (ASRCT) on the stability of a high pressure ratio centrifugal compressor by experimental testing and numerical simulation. Firstly, the performance of the compressor with/without SRCT is tested is measured investigate the influence of flow distortion on the stability of compressor as well as the numerical method validation. Then detailed flow field investigation is conducted by experimental measurement and the numerical method to unveil the reasons for stability enhancement by symmetrical/asymmetrical SRCT. Results show that static pressure distortion at impeller outlet caused by the volute can make passages be confronted with flow distortion less stable than others because of their larger positive slope of T-S pressure ratio performance at small flow rate. SRCT can depress the flow distortion and reduce the slope by non-uniform recycling flow rate at impeller inlet. Moreover, ASRCT can redistribute the recycling flow in circumferential direction according to the asymmetric geometries. When the largest recycling flow rate is imposed on the passage near the distorted static pressure, the slope will be the most effectively reduced. Therefore, the stability is effectively enhanced by the optimized recycling flow device.

Aerodynamics of a Transonic Centrifugal Compressor Impeller

2002 ◽  
Author(s):  
Seiichi Ibaraki ◽  
Tetsuya Matsuo ◽  
Hiroshi Kuma ◽  
Kunio Sumida ◽  
Toru Suita
Keyword(s):  
Shock Wave ◽  
High Pressure ◽  
Centrifugal Compressor ◽  
High Efficiency ◽  
Pressure Ratio ◽  
Flow Analysis ◽  
Flow Distortion ◽  
Flow Measurements ◽  
High Pressure Ratio ◽  
Compressor Impeller

High pressure ratio centrifugal compressors are applied to turbochargers and turboshaft engines because of their small dimensions, high efficiency and wide operating range. Such a high pressure ratio centrifugal compressor has a transonic inlet condition accompanied with a shock wave in the inducer portion. It is generally said that extra losses are generated by interaction of the shock wave and the boundary layers on the blade surface. To improve the performance of high pressure ratio centrifugal compressor it is necessary to understand the flow phenomena. Although some research works on transonic impeller flow have been published, some unknown flow physics are still remaining. The authors designed a transonic impeller, with an inlet Mach number is about 1.3, and conducted detailed flow measurements by using Laser Doppler Velocimetry (LDV). In the result the interaction between the shock wave and tip leakage vortex at the inducer and flow distortion at the downstream of inducer were observed. The interaction of the boundary layer and the shock wave was not observed. Also computational flow analysis were conducted and compared with experimental results.

Experiments With a Model Free Rotating Vaneless Diffuser

1975 ◽  
Vol 97 (2) ◽  
pp. 231-242 ◽  
Author(s):  
C. Rodgers ◽  
H. Mnew
Keyword(s):  
Static Pressure ◽  
Experimental Testing ◽  
Pressure Ratio ◽  
Velocity Ratio ◽  
Tangential Velocity ◽  
Vaneless Diffuser ◽  
Vaned Diffuser ◽  
High Pressure Ratio ◽  
Model Free ◽  
Inner Diameter

Experimental testing of a model free-rotating vaneless diffuser, for application to high pressure ratio single-stage centrifugal compressors, was conducted to determine diffuser performance under braked and free rotating conditions at entry Mach numbers up to unity. The experimental test rig comprised a swirl generating nozzle upstream of the model vaneless diffuser rotor with an outer-to-inner diameter ratio of 1.3. Additional downstream diffusion was completed with stationary vaneless and vaned diffuser inserts. A significant improvement in diffuser performance was achieved under free-rotating conditions even though large wakes generated by upstream stationary swirl nozzles were present. Overall static pressure recovery for the complete diffusion system increased approximately 20 percent at free-rotating conditions corresponding to a tangential velocity ratio (diffuser rotor/incident stream) of 0.43.

scholarly journals Experimental Investigations of a Recently Developed Supersonic Compressor Stage (Rotor and Stage Performance)

1974 ◽  
Author(s):  
H. Simon ◽  
D. Bohn
Keyword(s):  
Static Pressure ◽  
Pressure Ratio ◽  
Pressure Rise ◽  
Strong Shock ◽  
Wall Pressure ◽  
Experimental Investigations ◽  
Pressure Measurements ◽  
Compressor Stage ◽  
Pressure Transducers ◽  
Rotor Type

The experimental investigations of a recently developed supersonic compressor stage working with a strong shock wave both in the rotor and stator are described. The shock in the inlet area of the rotor is stabilized by the geometry of the rotor blade channel, whereas the position of shocks in the stator is controlled by the back pressure. Due to this, the static pressure rise is distributed to the rotor and stator avoiding a higher loading of the stator. In the first part the performance of the rotor alone has been investigated. The conducted probe and static wall pressure measurements allowed a detailed analysis of the flow through the rotor. The determined performance characteristics of the rotor show the peculiarities of the rotor at different speeds and throttle positions. In addition to the static wall pressure measurements the nonsteady pressure distributions have been measured at the casing wall by piezoelectric pressure transducers. Since this rotor (type ②) has been designed with the same relative inlet Mach number and turning as the previously investigated supersonic rotor (type ①), a direct comparison of these rotors can be made. In the second part of these investigations the rotor of type ② has been combined with a tandem cascade as a stator, to investigate the supersonic compressor stage. With heavy throttling a static pressure ratio of 3,5 (p3/p1) has been achieved. The evaluation of the probe measurements allowed a better estimation of the overall performance of this supersonic compressor stage.

The Experimental Study of Matching Between Centrifugal Compressor Impeller and Diffuser

1999 ◽  
Vol 121 (1) ◽  
pp. 113-118 ◽  
Author(s):  
H. Tamaki ◽  
H. Nakao ◽  
M. Saito
Keyword(s):  
Flow Rate ◽  
Centrifugal Compressor ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Design Flow ◽  
Negative Slope ◽  
Compressor Stage ◽  
Rate Analysis ◽  
System Pressure ◽  
Compressor Impeller

The centrifugal compressor for a marine use turbocharger with its design pressure ratio of 3.2 was tested with a vaneless diffuser and various vaned diffusers. Vaned diffusers were chosen to cover impeller operating range as broad as possible. The analysis of the static pressure ratio in the impeller and the diffusing system, consisting of the diffuser and scroll, showed that there were four possible combinations of characteristics of impeller pressure ratio and diffusing system pressure ratio, The flow rate, QP, where the impeller achieved maximum static pressure ratio, was surge flow rate of the centrifugal compressor determined by the critical flow rate. In order to operate the compressor at a rate lower than QP, the diffusing system, whose pressure recovery factor was steep negative slope near QP, was needed. When the diffuser throat area was less than a certain value, the compressor efficiency deteriorated; however, the compressor stage pressure ratio was almost constant. In this study, by reducing the diffuser throat area, the compressor could be operated at a flow rate less than 40 percent of its design flow rate. Analysis of the pressure ratio in the impeller and diffusing systems at design and off-design speeds showed that the irregularities in surge line occurred when the component that controlled the negative slope on the compressor stage pressure ratio changed.

Eliminating Static Pressure Distortion by a Large Cut Back Tongue Volute

2006 ◽  
Author(s):  
C. Xu ◽  
R. S. Amano
Keyword(s):  
High Pressure ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Experimental Studies ◽  
Flow Characteristics ◽  
Flow Structures ◽  
High Pressure Ratio ◽  
Compressor Performance ◽  
Stationary Frame ◽  
Good Agreement

This paper presents a physical solution by eliminating static pressure distortions of impeller exit due to a volute in a centrifugal compressor. The numerical and experimental studies on the circumferential distortion flow characteristics inside the stationary frame of a high-pressure ratio compressor with a large cut back tongue volute. The detailed flow structures and pressure distortions development inside the stationary components are discussed. The numerical results were demonstrated to be in good agreement with the experiments. The volute and diffuser interactions at design and off-design conditions were found to be much smaller for the large cut back volute in comparison with the reported from literature. The study indicated that the large cut back tongue volute design not only benefits the compressor performance but also reduces the impeller exit static pressure non-uniformity caused by discharge volute.

scholarly journals Stability Improvement of High-Pressure-Ratio Turbocharger Centrifugal Compressor by Asymmetric Flow Control—Part I: Non-Axisymmetrical Flow in Centrifugal Compressor

2012 ◽  
Vol 135 (2) ◽  
Author(s):  
Mingyang Yang ◽  
Xinqian Zheng ◽  
Yangjun Zhang ◽  
Takahiro Bamba ◽  
Hideaki Tamaki ◽  
...  
Keyword(s):  
High Pressure ◽  
Phase Shift ◽  
Flow Control ◽  
Centrifugal Compressor ◽  
Rotational Speed ◽  
Static Pressure ◽  
Control Method ◽  
Pressure Ratio ◽  
Asymmetric Flow ◽  
High Pressure Ratio

This is Part I of a two-part paper documenting the development of a novel asymmetric flow control method to improve the stability of a high-pressure-ratio turbocharger centrifugal compressor. Part I focuses on the nonaxisymmetrical flow in a centrifugal compressor induced by the nonaxisymmetrical geometry of the volute while Part II describes the development of an asymmetric flow control method to avoid the stall on the basis of the characteristic of nonaxisymmetrical flow. To understand the asymmetries, experimental measurements and corresponding numerical simulation were carried out. The static pressure was measured by probes at different circumferential and stream-wise positions to gain insights about the asymmetries. The experimental results show that there is an evident nonaxisymmetrical flow pattern throughout the compressor due to the asymmetric geometry of the overhung volute. The static pressure field in the diffuser is distorted at approximately 90 deg in the rotational direction of the volute tongue throughout the diffuser. The magnitude of this distortion slightly varies with the rotational speed. The magnitude of the static pressure distortion in the impeller is a function of the rotational speed. There is a significant phase shift between the static pressure distributions at the leading edge of the splitter blades and the impeller outlet. The numerical steady state simulation neglects the aforementioned unsteady effects found in the experiments and cannot predict the phase shift, however, a detailed asymmetric flow field structure is obviously obtained.

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