Throughflow Calculations for Transonic Axial Flow Turbines

1978 ◽  
Vol 100 (2) ◽  
pp. 212-218 ◽  
Author(s):  
J. D. Denton
Keyword(s):  
Transonic Flow ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Steam Turbines ◽  
Flow Measurements ◽  
High Pressure Ratio ◽  
Streamline Curvature ◽  
Flow Through ◽  
Good Agreement

The development of a streamline curvature throughflow program to predict the flow through the low pressure stages of large steam turbines is described. The program can also be used for gas turbines. Difficulties encountered in dealing with transonic flow and multiple high pressure ratio stages are discussed. Comparisons of the predictions of the program with flow measurements in steam and gas turbines show reasonably good agreement with most of the discrepancies being attributable to errors in the empirical data input to the program.

Influence of Open and Closed Shroud Cavities on the Flowfield in a 2-Stage Turbine With Shrouded Blades

2003 ◽  
Author(s):  
Dieter E. Bohn ◽  
Ingo Balkowski ◽  
Hongwei Ma ◽  
Christian Tu¨mmers ◽  
Michael Sell
Keyword(s):  
Flow Field ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Leading Edge ◽  
Operating Conditions ◽  
Complex Flow ◽  
Multi Stage ◽  
Flow Through ◽  
Insight Into

An important goal of the development of turbine bladings is to increase the efficiency for an optimized use of energy resources. This necessitates the most possible insight into the complex flow phenomena in multi-stage turbine bladings. This paper presents a combined numerical and experimental investigation of the flow field in a 2-stage axial turbine with shrouded blades, where the axial gap between the shroud and the endwall is varied between 1mm (closed cavities) and 5 mm (opened cavities). In the experimental setup at the Institute of Steam and Gas Turbines, Aachen University, the turbine is operated at a low pressure ratio of 1.4 with an inlet pressure of 3.2 bar. The rotating speed is adjusted by a water brake, which is integrated into a swing frame running in hydrostatic bearings. The rotor power dissipates in the water brake, which enables a very accurate angular momentum determination. The mass flow is measured through a calibrated nozzle installed upstream of the turbine inlet at an accuracy of better than 1%, from which stage efficiencies can be derived. For both geometric configurations (open and closed shroud cavities), the flow field at both inlet and outlet is measured using 5-hole probes as well as temperature probes at three operating conditions. The test rig is especially designed to investigate the influence of the cavity size. Therefore, the radial gaps between shroud and casing is held near zero in order to prevent an axial flow through the cavities. The experimental results are used as boundary conditions for corresponding numerical multi-stage calculations of the 3D flow through the 2-stage turbine, using the highly accurate steady Navier-Stokes inhouse computer code, CHT-Flow. The flow field measurements and the numerical simulations give deeper insight into some of the cavity-related flow field phenomena. The measurement results as well as the simulations indicate that the stator leading edge has little influence on the inlet flow field. The flow through the shroud cavities has a significant influence on the field and therefore on the machine’s performance.

Design Study of an Advanced Concept Simple Gas Turbine for Possible Use in Low Emission Automobiles

1972 ◽  
Author(s):  
H. C. Eatock ◽  
M. D. Stoten
Keyword(s):  
High Pressure ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Simple Cycle ◽  
Turbine Engines ◽  
Nox Emissions ◽  
Preliminary Design ◽  
Gas Turbine Engines ◽  
High Pressure Ratio

United Aircraft Corporation studied the potential costs of various possible gas turbine engines which might be used to reduce automobile exhaust emissions. As part of that study, United Aircraft of Canada undertook the preliminary design and performance analysis of high-pressure-ratio nonregenerated (simple cycle) gas turbine engines. For the first time, high levels of single-stage component efficiency are available extending from a pressure ratio less than 4 up to 10 or 12 to 1. As a result, the study showed that the simple-cycle engine may provide satisfactory running costs with significantly lower manufacturing costs and NOx emissions than a regenerated engine. In this paper some features of the preliminary design of both single-shaft and a free power turbine version of this engine are examined. The major component technology assumptions, in particular the high pressure ratio centrifugal compressor, employed for performance extrapolation are explained and compared with current technology. The potential low NOx emissions of the simple-cycle gas turbine compared to regenerative or recuperative gas turbines is discussed. Finally, some of the problems which might be encountered in using this totally different power plant for the conventional automobile are identified.

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 HTGR Helium Compressor Analysis Tool Based on Newton-Raphson Numerical Application to Throughflow Analysis

2006 ◽  
Author(s):  
Ji Hwan Kim ◽  
Hyeun Min Kim ◽  
Hee Cheon No
Keyword(s):  
Axial Flow ◽  
Meridional Flow ◽  
Analysis Tool ◽  
Numerical Application ◽  
Good Convergence ◽  
Streamline Curvature ◽  
Performance Map ◽  
Compressor Performance ◽  
Newton Raphson ◽  
Good Agreement

This study describes the development of a computer program for analyzing the off-design performance of axial flow helium compressors, which is one of the major concerns for the power conversion system of a high temperature gas-cooled reactor (HTGR). The compressor performance has been predicted by the aerodynamic analysis of meridional flow with allowances for losses. The governing equations have been derived from Euler turbomachine equation and the streamline curvature method, and then they have been merged into linearized equations based on the Newton-Raphson numerical method. The effect of viscosity is considered by empirical correlations to introduce entropy rises caused by primary loss sources. Use of the method has been illustrated by applying it to a 20-stage helium compressor of the GTHTR300 plant. As a result, the flow throughout the stages of the compressor has been predicted and the compressor characteristics have been also investigated according to the design specification. The program results show much better stability and good convergence with respect to other throughflow methods, and good agreement with the compressor performance map provided by JAEA.

Health Monitoring of Variable Geometry Gas Turbines for the Canadian Navy

1989 ◽  
Vol 111 (2) ◽  
pp. 244-250 ◽  
Author(s):  
D. E. Muir ◽  
H. I. H. Saravanamuttoo ◽  
D. J. Marshall
Keyword(s):  
Health Monitoring ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Engine Performance ◽  
Variable Geometry ◽  
Axial Compressors ◽  
High Pressure Ratio ◽  
Department Of National Defence ◽  
Canadian Navy ◽  
General Method

The Canadian Department of National Defence has identified a need for improved Engine Health Monitoring procedures for the new Canadian Patrol Frigate (CPF). The CPF propulsion system includes two General Electric LM2500 gas turbines, a high-pressure-ratio engine with multiple stages of compressor variable geometry. A general method for predicting the thermodynamic performance of variable geometry axial compressors has been developed. The new modeling technique is based on a meanline stage-stacking analysis and relies only on the limited performance data typically made available by engine manufacturers. The method has been applied to the LM2500-30 marine gas turbine and the variations in engine performance that can result from a malfunction of the variable geometry system in service have been estimated.

Deterioration of Compressor Performance Due to Tip Clearance of Centrifugal Impellers

1987 ◽  
Vol 109 (1) ◽  
pp. 55-61 ◽  
Author(s):  
Y. Senoo ◽  
M. Ishida
Keyword(s):  
High Pressure ◽  
Pressure Ratio ◽  
Tip Clearance ◽  
Slip Coefficient ◽  
Clearance Ratio ◽  
Flow Rates ◽  
High Pressure Ratio ◽  
Compressor Performance ◽  
Loss Efficiency ◽  
Good Agreement

The authors’ theory on the tip-clearance loss of centrifugal impellers is modified to include the variation of slip coefficient of the impeller due to the tip clearance, by deriving a rational relationship between two empirical parameters in the theory. In order to compare experimental data in the literature with prediction, examination was made regarding accuracy of available data and the way to select corresponding flow rates of a compressor with different values of tip clearance. Good agreement between data and prediction was observed. These examples demonstrate the following tendency regarding effects of various parameters on the tip clearance loss. Efficiency drop due to the tip clearance of high-pressure-ratio compressors is less than that of low-pressure-ratio compressors if the tip clearance ratio at the impeller exit is equal. The magnitude of clearance loss becomes smaller as the flow rate is reduced and also at a reduced shaft speed in cases of high-pressure-ratio compressors. The equations in the theory clearly show these tendencies.

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.

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.

scholarly journals Blade Vibration on Centrifugal Compressors: Fundamental Considerations and Initial Measurements

1985 ◽  
Author(s):  
U. Haupt ◽  
K. Bammert ◽  
M. Rautenberg
Keyword(s):  
Rotational Speed ◽  
Pressure Ratio ◽  
High Mass ◽  
Resonance Excitation ◽  
Centrifugal Compressors ◽  
Blade Vibration ◽  
Flow Measurements ◽  
High Pressure Ratio ◽  
Blade Thickness ◽  
Compressor Inlet

Blade vibration has to be considered in the design of high pressure ratio / high mass flow centrifugal compressors with increasing rotational speed values due to the reduced blade thickness. Results of a theoretical and experimental investigation concerning this problem are described. FE calculations of the stress distribution on the blade for the lower natural frequencies and various vibration tests at rest were carried out to investigate resonance and damping effects. This preparatory work was aimed at determining blade vibration behavior and acquiring fundamental experience for measurements on compressors in operation. Results of blade vibration measurements on compressors with a vaneless diffuser carried out with semiconductor strain gages and an 8-channel telemetry system are presented for constant rotational speed and for constant throttle valve position, and indicate considerable blade excitation during stall effects. Simultaneous flow measurements complete the investigation to determine the causes of blade vibration. For example, investigations were made of the extent of blade resonance excitation due to non-uniformity of the flow downstream of the impeller or due to flow disturbances caused by carrier blades for bearings in the compressor inlet and simulated by spoilers.

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