Approach to High-Performance Transonic Compressor Design

2004 ◽  
Vol 20 (1) ◽  
pp. 164-170 ◽  
Author(s):  
M. Oana ◽  
O. Kawamoto ◽  
H. Ohtani ◽  
Y. Yamamoto
Keyword(s):  
High Performance ◽  
Transonic Compressor ◽  
Compressor Design

Flow Field Unsteadiness in the Tip Region of a Transonic Compressor Rotor

1997 ◽  
Vol 119 (1) ◽  
pp. 122-128 ◽  
Author(s):  
S. L. Puterbaugh ◽  
W. W. Copenhaver
Keyword(s):  
Flow Field ◽  
High Performance ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Vortex Interaction ◽  
Tip Leakage Vortex ◽  
Tip Leakage ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Operating Points

An experimental investigation concerning tip flow field unsteadiness was performed for a high-performance, state-of-the-art transonic compressor rotor. Casing-mounted high frequency response pressure transducers were used to indicate both the ensemble averaged and time varying flow structure present in the tip region of the rotor at four different operating points at design speed. The ensemble averaged information revealed the shock structure as it evolved from a dual shock system at open throttle to an attached shock at peak efficiency to a detached orientation at near stall. Steady three-dimensional Navier Stokes analysis reveals the dominant flow structures in the tip region in support of the ensemble averaged measurements. A tip leakage vortex is evident at all operating points as regions of low static pressure and appears in the same location as the vortex found in the numerical solution. An unsteadiness parameter was calculated to quantify the unsteadiness in the tip cascade plane. In general, regions of peak unsteadiness appear near shocks and in the area interpreted as the shock-tip leakage vortex interaction. Local peaks of unsteadiness appear in mid-passage downstream of the shock-vortex interaction. Flow field features not evident in the ensemble averaged data are examined via a Navier-Stokes solution obtained at the near stall operating point.

Numerical Investigations on Partial Surge Initiated Inception and Stall Evolution in a Transonic Compressor

2017 ◽  
Author(s):  
Jiaguo Hu ◽  
Tianyu Pan ◽  
Wenqian Wu ◽  
Qiushi Li ◽  
Yifang Gong
Keyword(s):  
High Performance ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Rotating Stall ◽  
Rapid Decline ◽  
Second Phase ◽  
Stall Inception ◽  
Transonic Compressor ◽  
Transient Simulations ◽  
Two Factors

The instability has been the largest barrier of the high performance axial compressor in the past decades. Stall inception, which determines the route and the characteristics of instability evolution, has been extensively focused on. A new stall inception, “partial surge”, is discovered in the recent experiments. In this paper full-annulus transient simulations are performed to study the origin of partial surge initiated inception and explain the aerodynamic mechanism. The simulations show that the stall inception firstly occurs at the stator hub region, and then transfers to the rotor tip region. The compressor finally stalled by the tip region rotating stall. The stall evolution is in accord with the experiments. The stall evolution can be divided into three phases. In the first phase, the stator corner separation gradually merged with the adjacent passages, producing an annulus stall cell at the stator hub region. In the second phase, the total pressure rise of hub region emerges rapid decline due to the fast expansion of the annulus stall cell, but the tip region maintains its pressure rise. In the third phase, a new rotating stall cell appears at the rotor tip region, leading to the onset of fast drop of the tip region pressure rise. The stall cells transfer from hub region to the tip region is caused by two factors, the blockage of the hub region which transfers more load to the tip region, and the separation fluid fluctuations in stator domain which increase the circumferential non-uniformity in the rotor domain. High load and non-uniformity at the rotor tip region induce the final rotating stall.

Aeroelastic Investigation of a Transonic Research Compressor

2013 ◽  
Author(s):  
S. Leichtfuß ◽  
F. Holzinger ◽  
C. Brandstetter ◽  
F. Wartzek ◽  
H. P. Schiffer
Keyword(s):  
Sensitivity Study ◽  
Structural Damping ◽  
Structural Vibrations ◽  
Numerical Investigations ◽  
Transonic Compressor ◽  
Compressor Design ◽  
Blade Flutter ◽  
The One ◽  
German Aerospace ◽  
Highly Loaded

The trend in modern compressor design is towards higher stage loading and less structural damping, resulting in increased flutter risk. The understanding of the underlying aeroelastic effects, especially at highly loaded BLISK rotors, is small. This paper reports on the analysis of flutter phenomena in a modern transonic compressor. The geometry examined here is the one-and-a-half stage transonic research compressor operated by Technische Universität Darmstadt. High blade deflections recorded during throttling measurements point to an aerodynamic excitation. Therefore, numerical investigations are carried out using the CFD-Code TRACE developed at the German Aerospace Center (DLR). Simulations are compared to measured compressor speed lines to validate the steady state results. The open source Finite Element code CalculiX is used to simulate the rotor blade eigenmodes and -frequencies. The results are then used in time-linearized calculations to determine the onset of flutter. These calculations confirm that there is an aerodynamic excitation of the first torsional eigenmode and blade flutter is at risk. A sensitivity study is carried out to further investigate the aerodynamic conditions under which structural vibrations become unstable and to identify influencing factors.

Design of a High-Performance Axial Compressor for Utility Gas Turbine

1992 ◽  
Vol 114 (2) ◽  
pp. 277-286 ◽  
Author(s):  
A. Sehra ◽  
J. Bettner ◽  
A. Cohn
Keyword(s):  
Gas Turbine ◽  
High Performance ◽  
High Efficiency ◽  
Guide Vane ◽  
Design Flow ◽  
Design Parameters ◽  
Design System ◽  
Inlet Guide Vane ◽  
Variable Geometry ◽  
Compressor Design

An aerodynamic design study to configure a high-efficiency industrial-size gas turbine compressor is presented. This study was conducted using an advanced aircraft engine compressor design system. Starting with an initial configuration based on conventional design practice, compressor design parameters were progressively optimized. To improve the efficiency potential of this design further, several advanced design concepts (such as stator ends bends and velocity controlled airfoils) were introduced. The projected poly tropic efficiency of the final advanced concept compressor design having 19 axial stages was estimated at 92.8 percent, which is 2 to 3 percent higher than the current high-efficiency aircraft turbine engine compressors. The influence of variable geometry on the flow and efficiency (at design speed) was also investigated. Operation at 77 percent design flow with inlet guide vanes and front five variable stators is predicted to increase the compressor efficiency by 6 points as compared to conventional designs having only the inlet guide vane as variable geometry.

Tip Flow Fields in a Low Aspect Ratio Transonic Compressor

1995 ◽  
Author(s):  
P. Russler ◽  
D. Rabe ◽  
B. Cybyk ◽  
C. Hah
Keyword(s):  
Shock Wave ◽  
Aspect Ratio ◽  
Flow Field ◽  
High Frequency ◽  
High Performance ◽  
Flow Structures ◽  
Transonic Compressor ◽  
Low Aspect Ratio ◽  
Laser Data ◽  
Blade Tip

Experimental data and computational predictions are used to characterize the tip flow field of a high performance, low aspect ratio, transonic compressor. Flow structures near the first stage blade tip are monitored experimentally using two different data acquisition schemes. High frequency pressure and laser fringe anemometry data are used to experimentally define the tip flow structure. The high frequency pressure data were acquired with an array of pressure transducers mounted in the rotor casing. Laser data were acquired through a window in the same position. The transducer and laser data adequately define the shock structure at the tip. Both the movement of the shock wave in the blade passage during changes in compressor loading and the interaction between the shock wave and the tip leakage vortex are detected. Similar flow structures and compressor loading effects are numerically predicted using a three-dimensional Navier-Stokes algorithm. A fundamental understanding of the flow field at the blade tip is obtained using these three complementary methods.

A Preliminary Numerical CFD Analysis of Transonic Compressor Rotors When Subjected to Inlet Swirl Distortion

2011 ◽  
Author(s):  
Javier Castaneda ◽  
Ahad Mehdi ◽  
Domenico di Cugno ◽  
Vassilios Pachidis
Keyword(s):  
Vortex Core ◽  
High Performance ◽  
Preliminary Investigation ◽  
Angle Distribution ◽  
Transonic Compressor ◽  
Aspect Ratios ◽  
Swirl Angle ◽  
Inlet Swirl ◽  
System Life ◽  
The Impact

A preliminary investigation of a CFD capability to assess the impact of inlet swirl distortion on transonic compressor rotors has been carried out. In the late 1960s with the advent of turbo fan engines, industry and government agencies became increasingly aware of the inlet total pressure distortion problem. Since then, the inlet/engine compatibility assessment has become a significant issue within the propulsion system life cycle. Nowadays the development of high-performance military aircraft and UAV with maneuvers before unthinkable, entail considerable levels of inlet flow angularity. The importance of developing a rigorous methodology to understand the effect of inlet swirl distortion on turbomachinery has also become one of the major concerns of present day. NASA rotor 67 and 37 were selected for this investigation having different hub to tip radius and aspect ratios. The steady state CFD simulations were carried out on two types of inlet swirl distortion scenarios: Bulk swirl (both Rotor 37 and 67) and Ground vortex (only Rotor 67). A parametric study to define the swirl angle distribution for ground vortex cases was also accomplished. The non-dimensional ground clearance, wind conditions and core vortex location at the inlet/engine AIP were the parameters taken into account. The study carried out suggests that ground vortex core location and vortex rotational direction greatly affect the shift of the speedline. This emphasizes the importance of identifying the radial location of ingested vortex core at the AIP as the turbomachinery response differs depending on it. Similar shift in speedlines for the bulk swirl cases were also observed.

scholarly journals Effects of Unsteady Flow Interactions on the Performance of a Highly-Loaded Transonic Compressor Stage

2015 ◽  
Author(s):  
Chunill Hah
Keyword(s):  
Total Pressure ◽  
Suction Side ◽  
Pressure Side ◽  
Rotor Wake ◽  
Transonic Compressor ◽  
Eddy Simulation ◽  
Large Eddy ◽  
Compressor Design ◽  
Stator Blade ◽  
Total Temperature

The primary focus of this paper is to investigate the loss sources in an advanced GE transonic compressor design with high reaction and high stage loading. This advanced compressor has been investigated both experimentally and analytically in the past. The measured compressor efficiency is significantly lower than the efficiency calculated with various existing tools based on RANS and URANS. The general understanding is that some important flow physics in this modern compressor design are not represented in the current tools. To pinpoint the source of the efficiency miss, an advanced test with detailed flow traverse was performed for the front one and a half stage at the NASA Glenn Research Center. In the present paper, a Large Eddy Simulation (LES) is employed to determine whether a higher-fidelity simulation can pick up any additional flow physics that can explain past efficiency miss with RANS and URANS. The results from the Large Eddy Simulation were compared with the NASA test results and the GE interpretation of the test data. LES calculates lower total pressure and higher total temperature on the pressure side of the stator, resulting in large loss generation on the pressure side of the stator. On the other hand, existing tools based on the RANS and URANS do not calculate this high total temperature and low total pressure on the pressure side of the stator. The calculated loss through the stator from LES seems to match the measured data and the GE data interpretation. Detailed examination of the unsteady flow field from LES indicates that the accumulation of high loss near the pressure side of the stator is due to the interaction of the rotor wake with the stator blade. The strong rotor wake interacts quite differently with the pressure side of the stator than with the suction side of the stator blade. The concave curvature on the pressure side of the stator blade increases the mixing of the rotor wake with the pressure side boundary layer significantly. On the other hand, the convex curvature on the suction side of the stator blade decreases the mixing and the suction side blade boundary layer remains thin. The jet velocity in the rotor wake in the stator frame seems to magnify the curvature effect in addition to inviscid redistribution of wake fluid toward the pressure side of the blade.

Supersonic Compression Stage Design and Test Results

2004 ◽  
Author(s):  
Shawn P. Lawlor ◽  
John B. Hinkey ◽  
Steven G. Mackin ◽  
Scott Henderson ◽  
Jonathan Bucher ◽  
...  
Keyword(s):  
Power Systems ◽  
High Performance ◽  
High Efficiency ◽  
Pressure Ratio ◽  
Test Results ◽  
Stage Design ◽  
Concept System ◽  
Operational Characteristics ◽  
Compressor Design ◽  
Machinery Design

Ramgen Power Systems, Inc. (RPS) is developing a family of high performance supersonic compressors that combine many of the aspects of shock compression systems commonly used in supersonic flight inlet design with turbo-machinery design practices employed in conventional axial and centrifugal compressor design. The result is a high efficiency compressor that is capable of single stage pressure ratios in excess of those available in existing axial or centrifugal compressors. A variety of design configurations for land-based compressors utilizing this system have been explored. A proof-of-concept system has been designed to demonstrate the basic operational characteristics of this family of compressors when operating on air. The test unit was designed to process ~1.43 kg/s and produce a pressure ratio across the supersonic rotor of 2.25:1. The theory of operation of this system will be reviewed along with selected results from initial performance tests.

A Transonic Compressor Design Methodology Including the Influence of 3D Passage Shock Waves

1999 ◽  
Author(s):  
X. Gui ◽  
S. Zhou
Keyword(s):  
Shock Waves ◽  
Design Problem ◽  
Design Methodology ◽  
Axial Flow ◽  
Structure Model ◽  
Loss Model ◽  
Transonic Compressor ◽  
Compressor Design ◽  
Basic Equations ◽  
Shock Loss

Throughflow theory, as a design problem, is extensively used in the design of transonic axial flow compressors and fans. The losses caused by a simplified passage shock named shock loss model were taken into account in the design of transonic stages. However, the influences of shock waves on flow fields are still under development for transonic design. This problem is studied in the present paper. The basic equations of the throughflow method are re-derived under the concept of discontinuous passage averaging, and then, the distribution shock forces acied on the averaged throughflow fields are obtained. To evaluate these effects on averaged throughflow fields, a passage shock structure model is developed based on the previous models. Finally, the shock influences are assessed and compared in the present paper. It is demonstrated that the distribution shock forces have dramatically effects on the results of design.

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