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.

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.

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.

Design and Test of a Novel Highly-Loaded Compressor

2019 ◽  
Author(s):  
Chengwu Yang ◽  
Ge Han ◽  
Shengfeng Zhao ◽  
Xingen Lu ◽  
Yanfeng Zhang ◽  
...  
Keyword(s):  
Pressure Ratio ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Flow Angle ◽  
Stall Margin ◽  
High Pressure Ratio ◽  
Numerical Predictions ◽  
Highly Loaded

Abstract The blades of rear stages in small size core compressors are reduced to shorter than 20 mm or even less due to overall high pressure ratio. The growing of tip clearance-to-blade height ratio of the rear stages enhance the leakage flow and increase the possibility of a strong clearance sensitivity, thus limiting the compressor efficiency and stability. A new concept of compressor, namely diffuser passage compressor (DP), for small size core compressors was introduced. The design aims at making the compressors robust to tip clearance leakage flow by reducing pressure difference between pressure and suction surfaces. To validate the concept, the second stage of a two-stage highly loaded axial compressor was designed with DP rotor according to a diffuser map. The diffuser passage stage has the same inlet condition and loading as the conventional compressor (CNV) stage, of which the work coefficient is around 0.37. The predicted performance and flow field of the DP were compared with the conventional axial compressor in detail. The rig testing was supplemented with the numerical predictions. Results reveal that the throttle characteristic of DP indicates higher pressure rise and the loss reduction in tip clearance is mainly responsible for the performance improvement. For the compressor with DP, the pressure and flow angle are more uniform on exit plane. What’s more, the rotor with diffused passage reveals more robust than the conventional rotor at double clearance gap. Furthermore, the experimental data indicate that DP presents higher pressure rise at design and part speeds. At design speed, the stall margin was extended by 7.25%. Moreover, peak adiabatic efficiency of DP is also higher than that of CNV by about 0.7%.

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.

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.

Aerodynamic Design and Analysis of a High Pressure Ratio Aspirated Compressor Stage

2000 ◽  
Author(s):  
Ali A. Merchant ◽  
Mark Drela ◽  
Jack L. Kerrebrock ◽  
John J. Adamczyk ◽  
Mark Celestina
Keyword(s):  
Boundary Layer ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Adverse Pressure Gradient ◽  
Compressor Stage ◽  
High Pressure Ratio ◽  
Total Pressure Ratio ◽  
Isentropic Efficiency

The pressure ratio of axial compressor stages can be significantly increased by controlling the development of blade and endwall boundary layers in regions of adverse pressure gradient by means of boundary layer suction. This concept is validated and demonstrated through the design and analysis of a unique aspirated compressor stage which achieves a total pressure ratio of 3.5 at a tip speed of 1500 ft/s. The aspirated stage was designed using an axisymmetric through-flow code coupled with a quasi three-dimensional cascade plane code with inverse design capability. Validation of the completed design was carried out with three-dimensional Navier-Stokes calculations. Spanwise slots were used on the rotor and stator suction surfaces to bleed the boundary layer with a total suction requirement of 4% of the inlet mass flow. Additional bleed of 3% was also required on the hub and shroud near shock impingement locations. A three-dimensional viscous evaluation of the design showed good agreement with the quasi three-dimensional design intent, except in the endwall regions. The three-dimensional viscous analysis predicted a mass averaged total pressure ratio of 3.7 at an isentropic efficiency of 93% for the rotor, and a mass averaged total pressure ratio of 3.4 at an isentropic efficiency of 86% for the stage.

scholarly journals Design Methodology for Splittered Axial Compressor Rotors

1990 ◽  
Author(s):  
K.-L. Tzuoo ◽  
S. S. Hingorani ◽  
A. K. Sehra
Keyword(s):  
High Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Meridional Flow ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Viscous Effects ◽  
High Pressure Ratio ◽  
Rotor Design ◽  
Work Distribution

Recent trend toward lightweight, compact compression systems for advanced aircraft gas turbine engines has created a need for very high pressure ratio fan and compressor stages. One way of achieving pressure ratio in excess of 3:1 in an axial blade row is to introduce splitters (partial vanes) between the principal blades, a concept pioneered by Wennerstrom during early 70s for application in a 3:1 pressure ratio single axial stage. This paper presents an advanced methodology for high pressure ratio splittered rotor design. The methodology centers around combining a meridional flow calculation, an arbitrary meanline blade generation procedure, and 3-D inviscid and viscous analyses. Methods for specifying work distribution, solidity, loss, and deviation distributions, as well as the airfoil generation and splitter vane placement are discussed in detail. Importance of 3-D viscous effects along with results from a 3-D viscous calculation for Wennerstrom’s splittered rotor are also presented.

A Comparison of the Predicted and Measured Performance of High Pressure Ratio Centrifugal Compressor Diffusers

1972 ◽  
Author(s):  
D. P. Kenny
Keyword(s):  
Boundary Layer ◽  
High Pressure ◽  
Mach Number ◽  
Centrifugal Compressor ◽  
Pressure Ratio ◽  
Layer Growth ◽  
High Pressure Ratio ◽  
Wall Boundary Layer ◽  
Exit Mach Number ◽  
Boundary Layer Growth

A novel analysis of the hub and shroud wall boundary layer growth through the diffusing system of a centrifugal compressor is proposed to model the physical processes. It is shown that the diffuser throat blockage and total pressure loss characteristics can be accurately predicted for a 6:1 PR stage. The static pressure effectiveness and stalling limit are successfully predicted qualitatively, but are underestimated and overestimated by 14 and 12 percent respectively. It is argued that diffuser performance is largely controlled by the combined effect of the boundary layer conditions on the hub and shroud walls at impeller exit and the diffusion required to the diffuser throat. For this reason, it is contended that, for best performance at high pressure ratio (≃ 12:1), impeller exit Mach number must be minimized by employing zero to negative prewhirl at impeller entry which in turn maximizes impeller entry shroud relative Mach number. Performance maps are presented for a single-stage centrifugal compressor based on this premise with specific speed = 90. At 15, 12 and 101 PR, 72, 75 and 76.8 percent efficiency, respectively, were attached at 2.6 lb/sec.

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