A Numerical Study on the Effect of Bleed System on Starting Ability and Flow Performance of Rampressor Inlet

2014 ◽  
Vol 136 (12) ◽  
Author(s):  
Weijia Kang ◽  
Zhansheng Liu ◽  
Yu Wang ◽  
Yanyang Dong ◽  
Yong Sun
Keyword(s):  
Boundary Layer ◽  
Power Systems ◽  
Numerical Study ◽  
Pressure Ratio ◽  
Work Condition ◽  
High Pressure Ratio ◽  
Compressor Rotor ◽  
Improvement Effect ◽  
Negative Effect ◽  
Computational Fluid Dynamics Cfd

A unique supersonic compressor rotor with high pressure ratio, termed the Rampressor, is presented by Ramgen Power Systems, Inc., (RPS). Based on the models of Rampressor inlet, the inlet flow field with bleed system is numerically studied. Validation of the employed computational fluid dynamics (CFD) scheme is provided through test cases. The effects of boundary layer bleed location and bleed amount on Rampressor rotor inlet start and flow performance are analyzed. The results indicate that the boundary layer bleed has a significant effect for start and flow performance of Rampressor inlet. Boundary layer bleed technique has been applied to eliminate the emerging flow separation zone for enhancing Rampressor rotor inlet performance and enlarging its stable working range. The starting ability and flow performance of Rampressor inlet are efficiently improved by bleeding system, but the improvement effect is different for Rampressor inlet with different bleed location. Along the position of bleeding system moves forward, the range of Rampressor inlet normal work rotation speed is enlarged. The flow performance of Rampressor inlet improves obviously with the increment of bleed flow rate, and exit stability of Rampressor inlet enhances. And in the same back pressure work condition of Rampressor inlet, bleed system has been shown to be effective that exit stability of Rampressor inlet ameliorates, but the loss of compressed air from the bleed system has a negative effect on overall Rampressor inlet efficiency.

Measured and Predicted Performance of a High Pressure Ratio Supersonic Compressor Rotor

2008 ◽  
Author(s):  
Allan D. Grosvenor ◽  
David A. Taylor ◽  
Jonathan R. Bucher ◽  
Michael J. Aarnio ◽  
Paul M. Brown ◽  
...  
Keyword(s):  
Boundary Layer ◽  
High Pressure ◽  
Carbon Capture ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Compression System ◽  
High Pressure Ratio ◽  
Compressor Rotor ◽  
Numerical Predictions ◽  
Shock System

The testing of an 8:1 pressure ratio supersonic single axial compressor rotor referred to as Rampressor-2 is described. Design of this shockwave compression system is based on principles employed for supersonic intakes consisting of a multi-shock compression system and boundary layer treatment. The rotor consists of three blade passages within which the shock system is produced by a ramp, throat and diffuser contoured on the hub. The technology has been previously demonstrated in a 2.3:1 pressure ratio experimental test compressor (Rampressor-1). Measured performance is compared with numerical predictions. Further developments to improve Rampressor performance are discussed, and the appropriateness of this technology for Carbon Capture & Sequestration and LNG applications is highlighted.

scholarly journals A Numerical Study for Flow Excitation and Performance of Rampressor Inlet considering Rotor Motion

2014 ◽  
Vol 2014 ◽  
pp. 1-16 ◽  
Author(s):  
Weijia Kang ◽  
Zhansheng Liu ◽  
Jiangbo Lu ◽  
Yu Wang ◽  
Yanyang Dong
Keyword(s):  
Power Systems ◽  
Fundamental Frequency ◽  
Numerical Study ◽  
Pressure Ratio ◽  
Frequency Component ◽  
Inlet Flow ◽  
Compressor Rotor ◽  
Frequency Components ◽  
Subsonic Diffuser ◽  
And Performance

A unique supersonic compressor rotor with high pressure ratio, termed the Rampressor, is presented by Ramgen Power Systems, Inc. (RPS). In order to obtain the excitation characteristic and performance of Rampressor inlet flow field under external excitation, compression inlet flow of Rampressor is studied with considering Rampressor rotor whirling. Flow excitation characteristics and performance of Rampressor inlet are analyzed under different frequency and amplitude of Rampressor rotor whirling. The results indicate that the rotor whirling has a significant effect for flow excitation characteristics and performance of Rampressor inlet. The effect of rotor whirling on the different inlet location excitation has a definite phase difference. Inlet excitation becomes more complex along with the inlet flow path. More frequency components appear in the excitation spectrum of Rampressor inlet with considering Rampressor rotor whirling. The main frequency component is the fundamental frequency, which is caused by the rotor whirling. Besides the fundamental frequency, the double frequency components are generated due to the coupling between inlet compression flow of Rampressor rotor and rotor whirling, especially in the subsonic diffuser of Rampressor rotor inlet. With the increment of rotor whirling frequency and whirling amplitude, the complexity of Rampressor inlet excitation increases, and the stability of Rampressor inlet performance deteriorates.

scholarly journals Effect of Partial Span Aspiration on the Performance of a Transonic Axial Compressor Rotor: A Numerical Study

2015 ◽  
Vol 2015 ◽  
pp. 1-13
Author(s):  
Vijaykumar Jain ◽  
Quamber H. Nagpurwala ◽  
Abdul Nassar
Keyword(s):  
Boundary Layer ◽  
Numerical Study ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Suction Side ◽  
Surface Boundary ◽  
Suction Surface ◽  
Stall Margin ◽  
Compressor Rotor ◽  
Transonic Axial Compressor

Aspiration in an axial compressor is normally regarded as sucking out the low momentum boundary layer from blade suction surface, thus lowering the chances of flow separation and consequently that of stall under off-design operation. However, the suction mass flow does not take part in useful work and leads to loss of engine power output. This paper deals with a new concept of natural aspiration to energize blade suction surface boundary layer by injecting some fluid from pressure to suction side through a part span slot on the blade. The energized boundary layer has lesser tendency to separate, thus enhancing stall margin. Numerical simulations were carried out to study the effect of aspiration slot location and geometry on the performance and stall margin of a transonic axial compressor rotor. The computational results without aspiration slot were in fair agreement with the published experimental data. The modified rotor, with part span aspiration, showed ~3.2% improvement in stall margin at design rotational speed. The pressure ratio and efficiency of the aspirated rotor dropped by ~1.42% and ~2.0%, respectively, whereas the structural analysis did not indicate any adverse effect on the blade stress distribution in the presence of aspiration slot.

Design Methodology and Predicted Performance for a Supersonic Compressor Stage

2006 ◽  
Author(s):  
Allan D. Grosvenor ◽  
Paul M. Brown ◽  
Shawn P. Lawlor
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
High Speed ◽  
Pressure Ratio ◽  
Compression System ◽  
High Pressure Ratio ◽  
Separation Shock ◽  
Compressor Rotor ◽  
Boundary Layer Interaction ◽  
Shock System

Aspects of the aerodynamic design of a unique supersonic high pressure ratio compressor rotor, termed the Rampressor, are presented. The design of this shock wave compression system is based on principles employed in supersonic intake design with a multi-shock compression system and boundary layer treatment. One of the unique features of this configuration is the way these techniques have been applied to the design of a high-speed rotor, as opposed to a system designed for linear flight. The rotor consists of three blade-rows within which the shock system is produced by a ramp, throat, and diffuser on the hub. The technology has been previously demonstrated in a 2.3:1 pressure ratio experimental test compressor. The present study concentrates on applying the same techniques to achieve pressure ratios in the range of 8–10:1. Estimated performance is supported by mean-line and method of characteristics calculations, as well as 3D viscous Computational Fluid Dynamics (CFD) simulations. Validation of the employed CFD scheme is provided through test cases that represent the physics of boundary layers, diffusing flows and separation, shock wave / boundary layer interaction, and compressor aerodynamics. The study concentrates on the predicted effect of hub contour on the rotor shock system, and subsequent impact on compressor performance.

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.

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.

Experimental and Analytical Investigation of the Effects of Reynolds Number and Blade Surface Roughness on Multistage Axial Flow Compressors

1980 ◽  
Vol 102 (1) ◽  
pp. 5-12 ◽  
Author(s):  
A. Scha¨ffler
Keyword(s):  
Experimental Data ◽  
Boundary Layer ◽  
Reynolds Number ◽  
Critical Reynolds Number ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Blade Surface ◽  
Laminar Separation ◽  
High Pressure Ratio ◽  
Attached Flow

The general effect of Reynolds Number on axial flow compressors operating over a sufficiently wide range is described and illustrated by experimental data for four multistage axial compressors. The wide operating range of military aircraft engines leads in the back stages of high pressure ratio compression systems to three distinctly different regimes of operation, characterized by the boundary layer conditions of the cascade flow: • laminar separation, • turbulent attached flow with hydraulically smooth blade surface, • turbulent attached flow with hydraulically rough blade surface. Two “critical” Reynolds Numbers are defined, the “lower critical Reynolds Number” below which laminar separation occurs with a definite steepening of the efficiency/Reynolds Number relation and an “upper critical Reynolds Number” above which the blade surface behaves hydraulically rough, resulting in an efficiency independant of Reynolds Number. The permissible blade surface roughness for hydraulically smooth boundary layer conditions in modern high pressure ratio compression systems is derived from experimental data achieved with blades produced by grinding, electrochemical machining and forging. A correlation between the effect of technical roughness and sand type roughness is given. The potential loss of efficiency in the back end of compression systems due to excessive blade roughness is derived from experimental results. The repeatedly experienced different sensitivity of front and back stages towards laminar separation in the low Reynolds Number regime is explained by boundary layer calculations as a Mach Number effect on blade pressure distribution, i.e. transonic versus subsonic flow.

Assessment of Various Turbulence Models in a High Pressure Ratio Centrifugal Compressor With an Object Oriented CFD Code

2012 ◽  
Vol 134 (6) ◽  
Author(s):  
Luca Mangani ◽  
Ernesto Casartelli ◽  
Sebastiano Mauri
Keyword(s):  
Boundary Layer ◽  
High Pressure ◽  
Pressure Gradient ◽  
Centrifugal Compressor ◽  
Pressure Ratio ◽  
Turbulence Models ◽  
Object Oriented ◽  
Adverse Pressure Gradient ◽  
High Mass ◽  
High Pressure Ratio

The flow field in a high pressure ratio centrifugal compressor with a vaneless diffuser has been investigated numerically. The main goal is to assess the influence of various turbulence models suitable for internal flows with an adverse pressure gradient. The numerical analysis is performed with a 3D RANS in-house modified solver based on an object-oriented open-source library. According to previous studies from varying authors, the turbulence model is believed to be the key parameter for the discrepancy between experimental and numerical results, especially at high pressure ratios and high mass-flow. Particular care has been taken at the wall, where a detailed integration of the boundary layer has been applied. The results present different comparisons between the models and experimental data, showing the influence of using advanced turbulence models. This is done in order to capture the boundary layer behavior, especially in large adverse pressure gradient single stage machinery.

A Theoretical Investigation of Transonic Flows in Radial Compressor Diffusers

1983 ◽  
Vol 105 (3) ◽  
pp. 452-456
Author(s):  
H. O. Jeske ◽  
I. Teipel
Keyword(s):  
Boundary Layer ◽  
Boundary Layer Flow ◽  
Pressure Ratio ◽  
Numerical Procedure ◽  
Vaned Diffuser ◽  
High Pressure Ratio ◽  
Integral Methods ◽  
Boundary Layer Interaction ◽  
Flow Through ◽  
Layer Flow

The transonic flow in a diffuser of a centrifugal compressor with high pressure ratio has been analyzed by a numerical procedure. The method consists of an inviscid calculation of the pressure field in the vaned diffuser and of a determination of the boundary layer flow along the blades. The diffuser has been equipped with curved vanes, and only the flow through one channel is considered. The two-dimensional pressure distribution has been calculated by a time-dependent finite difference scheme. The boundary layer flow has been determined by different integral methods with special attention concerning the shock-boundary-layer interaction. Finally, the numerical results are compared with experiments, and the agreement is satisfactory.

Highly-Loaded Low-Reaction Boundary Layer Suction Axial Flow Compressor

2007 ◽  
Author(s):  
Songtao Wang ◽  
Xiaoqing Qiang ◽  
Weichun Lin ◽  
Guotai Feng ◽  
Zhongqi Wang
Keyword(s):  
Boundary Layer ◽  
High Pressure ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Design Concept ◽  
Axial Flow Compressor ◽  
Boundary Layer Suction ◽  
High Pressure Ratio ◽  
Flow Compressor ◽  
Highly Loaded

In order to design high pressure ratio and highly loaded axial flow compressor, a new design concept based on Highly-Loaded Low-Reaction and boundary layer suction was proposed in this paper. Then the concept’s characteristics were pointed out by comparing with the MIT’s boundary layer suction compressor. Also the application area of this design concept and its key technic were given out in this paper. Two applications were carried out in order to demonstrate the concept. The first application was to redesign a low speed duplication-stage axial flow compressor into a single stage. The second one was a feasibility analysis to decrease an 11 stage axial compressor’s stage count to 7 while not changing its aerodynamic performance. The analysis result showed that the new design concept is feasible and it can be used on high pressure stage of the aero-engine, compressor of ground gas turbine (except the transonic stage) and high total pressure ratio blower.

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