The Design, Performance and Analysis of a High Work Capacity Transonic Turbine

1985 ◽  
Vol 107 (4) ◽  
pp. 931-937 ◽  
Author(s):  
J. D. Bryce ◽  
M. R. Litchfield ◽  
N. P. Leversuch
Keyword(s):  
Low Cost ◽  
Pressure Ratio ◽  
Work Capacity ◽  
Tip Clearance ◽  
Total Efficiency ◽  
Point Pressure ◽  
Shaft Power ◽  
Transonic Turbine ◽  
Design And Testing ◽  
High Work

This paper describes the design and testing of a high work capacity single-stage transonic turbine of aerodynamic duty tailored to the requirements of driving the high-pressure core of a low cost turbofan engine. Aerodynamic loading was high for this duty (ΔH/U2 = 2.1) and a major objective in the design was the control of the resulting transonic flow to achieve good turbine performance. Practical and coolable blading was a design requirement. At the design point (pressure ratio = 4.48), a turbine total to total efficiency of 87.0 percent was measured—this being based on measured shaft power and a tip clearance of 1.4 percent of blade height. In addition, the turbine was comprehensively instrumented to allow measurement of aerofoil surface static pressures on both stator and rotor—the latter being expedited via a rotating scanivalve system. Downstream area traverses were also conducted. Analysis of these measurements indicates that the turbine operates at overall reaction levels lower than design but the rotor blade performs efficiently.

An Investigation of the Flow Structure in the Radial Inflow Transonic Turbine

2015 ◽  
Author(s):  
Cheng Zhu ◽  
Weilin Zhuge ◽  
Yangjun Zhang
Keyword(s):  
High Pressure ◽  
Secondary Flow ◽  
Flow Structure ◽  
Cross Sections ◽  
Low Cost ◽  
Pressure Ratio ◽  
Flow Characteristics ◽  
Horseshoe Vortex ◽  
High Pressure Ratio ◽  
Transonic Turbine

Radial inflow turbines which are an important component of a turbocharger consist essentially of a volute, a rotor and a diffuser. Vaneless volute turbines, which have reasonable performance and low cost, are the most widely used in turbochargers for automotive engines. In recent years the growing necessity of increasing specific output power of turbochargers has encouraged the design of high pressure ratio turbine stage. Two stage turbines, which can achieve the high pressure ratio require, are not suitable to for these applications due to volume and weight increases. The common design trend is thus to use single stage high pressure ratio radial transonic turbine. This paper describes numerical investigations of the flow fields in a radial inflow transonic turbine whose design pressure ratio is 4. The S-A turbulence model and Jameson’s center scheme have been applied in order to get good viscous resolution, accuracy and computing efficiency. Limiting streamlines on the wall surface as well as different flow characteristics, such as entropy generation of the cross sections, were evaluated, and detailed endwall flow and secondary flow structure were analyzed. The development of different vortex, especially the tip leakage vortex, vortex caused by the shock wave, passage vortex and horseshoe vortex were discussed. The results have shown that there is a great secondary flow feature and complicated vortex system in the high pressure ratio radial inflow transonic turbine.

scholarly journals Efficiency Scaling - Influence of Reynolds and Mach Number on Fan Performance

2021 ◽  
pp. 1-17
Author(s):  
Peter F. Pelz ◽  
Sebastian Saul ◽  
Johannes Brötz
Keyword(s):  
Mach Number ◽  
Standardized Test ◽  
Pressure Ratio ◽  
Reference Method ◽  
Operating Conditions ◽  
Tip Clearance ◽  
Scaling Method ◽  
Scaled Model ◽  
Type Size ◽  
Shaft Power

Abstract The efficiency, pressure ratio and shaft power of a fan depends on type, size, working medium and operating condition. For acceptance tests, a dissimilarity in Reynolds number, Mach number, relative roughness and relative blade tip clearance of the scaled model and prototype is unavoidable. Hence, the efficiency differs between model and prototype. This difference is quantified by scaling methods. This paper presents a validated and physics based, i. e. reliable scaling method for the efficiency, pressure ratio and shaft power of axial and centrifugal fans operating at subsonic conditions. The method is validated using test results gained on standardized test rigs for different fan types, sizes and operating conditions. For all scenarios the presented scaling method provides a much reduced scaling uncertainty compared to the reference method described in ISO 13348.

scholarly journals Design and Development of an Advanced Two-Stage Centrifugal Compressor

1994 ◽  
Author(s):  
D. L. Palmer ◽  
W. F. Waterman
Keyword(s):  
Centrifugal Compressor ◽  
High Performance ◽  
Low Cost ◽  
Pressure Ratio ◽  
Tip Clearance ◽  
Development Effort ◽  
Design And Development ◽  
Two Stage ◽  
Surge Margin ◽  
The Impact

This paper describes the aero-mechanical design and development of a 3.3 kg/sec (7.3 lb/sec), 14:1 pressure ratio two-stage centrifugal compressor which is used in the T800-LHT-800 helicopter engine. The design employs highly nonradial, splitter bladed impellers with swept leading edges and compact vaned diffusers to achieve high performance in a small and robust configuration. The development effort quantified the effects of impeller diffusion and passive inducer shroud bleed on surge margin as well as the effects of impeller loading on tip clearance sensitivity and the impact of sand erosion and shroud roughness on performance. The developed compressor exceeded its performance objectives with a minimum of 23-percent surge margin without variable geometry. The compressor provides a high performance, rugged, low-cost configuration ideally suited for helicopter applications.

Evaluation of a High Hub/Tip Ratio Centrifugal Compressor

1970 ◽  
Vol 92 (3) ◽  
pp. 419-428 ◽  
Author(s):  
F. G. Groh ◽  
G. M. Wood ◽  
R. S. Kulp ◽  
D. P. Kenny
Keyword(s):  
Centrifugal Compressor ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Tip Clearance ◽  
Total Efficiency ◽  
Compressor Stage ◽  
Vaneless Diffuser ◽  
Centrifugal Stage ◽  
Design Speed ◽  
Multistage Axial Compressor

A centrifugal compressor stage with an unusually high inlet hub/tip ratio of 0.87 was designed for a pressure ratio of 2.0 at a corrected mass flow of 2.45 lb per sec. The geometry was selected so that the centrifugal stage could replace several of the last stages of a multistage axial compressor. The stage was tested with two diffuser schemes. One diffuser consisted of a series of drilled conical pipes, whereas the other employed multirow vaned cascades. Sea level aerodynamic tests of the compressor stage with each diffuser showed a peak total-to-total efficiency at design speed of 83.8 percent for the pipe diffuser and 82.9 percent for the vaned cascade diffuser. Additional tests were conducted with a vaneless diffuser to determine effects of impeller discharge tip clearance and inlet prewhirl on impeller performance.

Design and Development of an Advanced Two-Stage Centrifugal Compressor

1995 ◽  
Vol 117 (2) ◽  
pp. 205-212 ◽  
Author(s):  
D. L. Palmer ◽  
W. F. Waterman
Keyword(s):  
Centrifugal Compressor ◽  
High Performance ◽  
Low Cost ◽  
Pressure Ratio ◽  
Tip Clearance ◽  
Development Effort ◽  
Design And Development ◽  
Two Stage ◽  
Surge Margin ◽  
The Impact

This paper describes the aeromechanical design and development of a 3.3 kg/s (7.3 lb/sec), 14:1 pressure ratio two-stage centrifugal compressor, which is used in the T800-LHT-800 helicopter engine. The design employs highly nonradial, splitter bladed impellers with swept leading edges and compact vaned diffusers to achieve high performance in a small and robust configuration. The development effort quantified the effects of impeller diffusion and passive inducer shroud bleed on surge margin as well as the effects of impeller loading on tip clearance sensitivity and the impact of sand erosion and shroud roughness on performance. The developed compressor exceeded its performance objectives with a minimum of 23 percent surge margin without variable geometry. The compressor provides a high-performance, rugged, low-cost configuration ideally suited for helicopter applications.

scholarly journals 5 MW Closed Cycle Gas Turbine

1984 ◽  
Author(s):  
John L. Mason ◽  
Anthony Pietsch ◽  
Theodore R. Wilson ◽  
Allen D. Harper
Keyword(s):  
Power Systems ◽  
Gas Turbine ◽  
Oil Recovery ◽  
Low Cost ◽  
Pressure Ratio ◽  
Commercial Application ◽  
Solid Fuels ◽  
Closed Cycle ◽  
Emission Standards ◽  
Fluidized Bed Combustor

A novel closed-cycle gas turbine power system is now under development by the GWF Power Systems Company for cogeneration applications. Nominally the system produces 5 megawatts (MW) of electric power and 80,000 lb/hr (36,287 kg/hr) of 1000 psig (6895 kPa) steam. The heat source is an atmospheric fluidized bed combustor (AFBC) capable of using low-cost solid fuels while meeting applicable emission standards. A simple, low-pressure ratio, single spool, turbomachine is utilized. This paper describes the system and related performance, as well as the development and test efforts now being conducted. The initial commercial application of the system will be for Enhanced Oil Recovery (EOR) of the heavy crudes produced in California.

Numerical Modeling of Heat Transfer and Pressure Losses for Uncooled Gas Turbine Blade Tip: Effect of Tip Clearance and Tip Geometry

2007 ◽  
Author(s):  
Lamyaa A. El-Gabry
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Gas Turbine ◽  
Computational Study ◽  
Numerical Models ◽  
Pressure Ratio ◽  
Tip Clearance ◽  
Pressure Losses ◽  
Blade Tip ◽  
Exit Mach Number

A computational study has been performed to predict the heat transfer distribution on the blade tip surface for a representative gas turbine first stage blade. CFD predictions of blade tip heat transfer are compared to test measurements taken in a linear cascade, when available. The blade geometry has an inlet Mach number of 0.3 and an exit Mach number of 0.75, pressure ratio of 1.5, exit Reynolds number based on axial chord of 2.57×106, and total turning of 110 deg. Three blade tip configurations were considered; they are flat tip, a full perimeter squealer, and an offset squealer where the rim is offset to the interior of the tip perimeter. These three tip geometries were modeled at three tip clearances of 1.25, 2.0, and 2.75% of blade span. The tip heat transfer results of the numerical models agree fairly well with the data and are comparable to other CFD predictions in the open literature.

Effects of Circumferential Nonuniform Tip Clearance on Flow Field and Performance of a Transonic Turbine

2021 ◽  
Author(s):  
Yun Zheng ◽  
Xiubo Jin ◽  
Hui Yang ◽  
Qingzhe Gao ◽  
Kang Xu
Keyword(s):  
Flow Field ◽  
Tip Clearance ◽  
Transonic Turbine ◽  
And Performance

The Design and Performance of a High Work Research Turbine

1996 ◽  
Vol 118 (4) ◽  
pp. 792-799 ◽  
Author(s):  
E. P. Vlasic ◽  
S. Girgis ◽  
S. H. Moustapha
Keyword(s):  
Pressure Ratio ◽  
Limit Load ◽  
Gas Generator ◽  
Part Load ◽  
Cold Flow ◽  
Turbine Efficiency ◽  
Load Pattern ◽  
Design Speed ◽  
And Performance ◽  
High Work

This paper describes the design and performance of a high work single-stage research turbine with a pressure ratio of 5.0, a stage loading of 2.2, and cooled stator and rotor. Tests were carried out in a cold flow rig and as part of a gas generator facility. The performance of the turbine was assessed, through measurements of reaction, rotor exit conditions and efficiency, with and without airfoil cooling. The measured cooled efficiency in the cold rig was 79.9 percent, which, after correcting for temperature and measuring plane location, matched reasonably well the efficiency of 81.5 percent in the gas generator test. The effect of cooling, as measured in the cold rig, was to reduce the turbine efficiency by 2.1 percent. A part-load turbine map was obtained at 100, 110, and 118 percent design speed and at 3.9, 5.0, and 6.0 pressure ratio. The influence of speed and the limit load pattern for transonic turbines are discussed. The effect of the downstream measuring distance on the calculated efficiency was determined using three different locations. An efficiency drop of 3.2 percent was measured between the rotor trailing edge plane and a distance four chords downstream.

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