A Feasibility Study of a High-Speed Combustor for Turbomachine Applications

1970 ◽  
Author(s):  
Irving Fruchtman
Keyword(s):  
Turbulent Mixing ◽  
High Speed ◽  
Pressure Ratio ◽  
Small Mass ◽  
Temperature Profiles ◽  
Flow Conditions ◽  
Finite Rate ◽  
High Pressure Ratio ◽  
Series Of Experiments ◽  
Analysis Design

The theoretical analysis, design, and experimental study of a high-speed combustion chamber are described. Such a burner may be used when the compressor outflow speed is so high that diffusion to the usual burner entrance conditions presents severe loss penalties. The study showed for a small mass flow-high pressure ratio turbomachine, that combined diffusor and combustor losses are minimum for a burner entrance Mach number of about 0.5. To design the burner a finite rate chemistry and turbulent mixing computer program was used; the combustor modeling and flame spread predictions are discussed. A series of experiments are described and burner pressure loss and temperature profiles are shown over a range of burner air-flow conditions.

scholarly journals The Design and Evaluation of a High Pressure Ratio Radial Turbine

1992 ◽  
Author(s):  
K. R. Pullen ◽  
N. C. Baines ◽  
S. H. Hill
Keyword(s):  
High Pressure ◽  
High Speed ◽  
Pressure Ratio ◽  
Performance Measurements ◽  
Turbine Engine ◽  
High Pressure Ratio ◽  
Turbine Efficiency ◽  
Wide Range ◽  
Dimensional Parameters ◽  
Very High

A single stage, high speed, high pressure ratio radial inflow turbine was designed for a single shaft gas turbine engine in the 200 kW power range. A model turbine has been tested in a cold rig facility with correct simulation of the important non-dimensional parameters. Performance measurements over a wide range of operation were made, together with extensive volute and exhaust traverses, so that gas velocities and incidence and deviation angles could be deduced. The turbine efficiency was lower than expected at all but the lowest speed. The rotor incidence and exit swirl angles, as obtained from the rig test data, were very similar to the design assumptions. However, evidence was found of a region of separation in the nozzle vane passages, presumably caused by a very high curvature in the endwall just upstream of the vane leading edges. The effects of such a separation are shown to be consistent with the observed performance.

Correlations Hidden in Compressor Maps

2011 ◽  
Author(s):  
Joachim Kurzke
Keyword(s):  
High Speed ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Operating Conditions ◽  
Specific Work ◽  
Gas Generator ◽  
High Pressure Ratio ◽  
Map Adaptation ◽  
Physical Phenomena ◽  
Corrected Flow

Realistic compressor maps are the key to high quality gas turbine performance calculations. When modeling the performance of an existing engine then these maps are usually not known and must be approximated by adapting maps from literature to either measured data or to other available information. There are many publications describing map adaptation processes, simple ones and more sophisticated physically based scaling rules. There are also reports about using statistics, genetic algorithms, neural networks and even morphing techniques for re-engineering compressor maps. This type of methods does not consider the laws of physics and consequently the generated maps are valid at best in the region in which they have been calibrated. This region is frequently very narrow, especially in case of gas generator compressors which run in steady state always on a single operating line. This paper describes which physical phenomena influence the shape of speed and efficiency lines in compressor maps. For machines operating at comparatively low speeds (so that the flow into each stage is subsonic), there is usually considerable range between choke and stall corrected flow. As the speed of the machine is increased the range narrows. For high-speed stages with supersonic relative flow into the rotor the efficiency maximum is where the speed line turns over from vertical to lower than maximum corrected flow. At this operating condition the shock is about to detach from the leading edge of the blades. The flow at a certain speed can also be limited by choking in the compressor exit guide vanes. For high pressure ratio single stage centrifugal compressors this is a normal case, but it can also happen with low pressure ratio multistage boosters of turbofan engines, for example. If the compressor chokes at the exit, then the specific work remains constant along the speed line while the overall pressure ratio varies and that generates a very specific shape of the efficiency contour lines in the map. Also in other parts of the map, the efficiency varies along speed lines in a systematic manner. Peculiar shapes of specific work and corrected torque lines can reveal physically impossibilities that are difficult to see in the standard compressor map pictures. Compressor maps generated without considering the inherent physical phenomena can easily result in misleading performance calculations if used at operating conditions outside of the region where they have been calibrated. Whatever map adaptation method is used: the maps created in such a way should be checked thoroughly for violations of the underlying laws of compressor physics.

Design and Test Results of a Ultra High Loaded Single Stage High Pressure Turbine

2013 ◽  
Author(s):  
Harjit S. Hura ◽  
Scott Carson ◽  
Rob Saeidi ◽  
Hyoun-Woo Shin ◽  
Paul Giel
Keyword(s):  
High Pressure ◽  
Total Pressure ◽  
High Speed ◽  
Pressure Ratio ◽  
Single Stage ◽  
Test Results ◽  
High Pressure Turbine ◽  
Technology Program ◽  
High Pressure Ratio ◽  
Expected Performance

This paper describes the engine and rig design, and test results of an ultra-highly loaded single stage high pressure turbine. In service aviation single stage HPTs typically operate at a total-to-total pressure ratio of less than 4.0. At higher pressure ratios or energy extraction the nozzle and blade both have regions of supersonic flow and shock structures which, if not mitigated, can result in a large loss in efficiency both in the turbine itself and due to interaction with the downstream component which may be a turbine center frame or a low pressure turbine. Extending the viability of the single stage HPT to higher pressure ratios is attractive as it enables a compact engine with less weight, and lower initial and maintenance costs as compared to a two stage HPT. The present work was performed as part of the NASA UEET (Ultra-Efficient Engine Technology) program from 2002 through 2005. The goal of the program was to design and rig test a cooled single stage HPT with a pressure ratio of 5.5 with an efficiency at least two points higher than the state of the art. Preliminary design tools and a design of experiments approach were used to design the flow path. Stage loading and through-flow were set at appropriate levels based on prior experience on high pressure ratio single stage turbines. Appropriate choices of blade aspect ratio, count, and reaction were made based on comparison with similar HPT designs. A low shock blading design approach was used to minimize the shock strength in the blade during design iterations. CFD calculations were made to assess performance. The HPT aerodynamics and cooling design was replicated and tested in a high speed rig at design point and off-design conditions. The turbine met or exceeded the expected performance level based on both steady state and radial/circumferential traverse data. High frequency dynamic total pressure measurements were made to understand the presence of unsteadiness that persists in the exhaust of a transonic turbine.

Performance of a Mixed Flow Turbocharger Turbine Under Pulsating Flow Conditions

1995 ◽  
Author(s):  
C. Arcoumanis ◽  
I. Hakeem ◽  
L. Khezzar ◽  
R. F. Martinez-Botas ◽  
N. C. Baines
Keyword(s):  
High Efficiency ◽  
Pressure Ratio ◽  
Velocity Ratio ◽  
Pulsating Flow ◽  
Flow Conditions ◽  
Mixed Flow ◽  
Engine Exhaust ◽  
High Pressure Ratio ◽  
Design Speed ◽  
Swallowing Capacity

The performance of a high pressure ratio (P.R.=2.9) mixed flow turbine for an automotive turbocharger has been investigated and the results revealed its better performance relative to a radial-inflow geometry under both steady and pulsating flow conditions. The advantages offered by the constant blade angle rotor allow better turbocharger-engine matching and maximization of the energy extracted from the pulsating engine exhaust gases. In particular, the mixed inlet blade geometry resulted in high efficiency at high expansion ratios where the engine-exhaust pulse energy is maximum. The efficiency characteristics of the mixed flow turbine under steady conditions were found to be fairly uniform when plotted against the velocity ratio, with a peak efficiency at the design speed of 0.75. The unsteady performance as indicated by the mass-averaged total-to-static efficiency and the swallowing capacity exhibited a departure from the quasi-steady assumption which is analysed and discussed.

Analysis of Instability Inception in High-Speed Multistage Axial-Flow Compressors

1997 ◽  
Vol 119 (4) ◽  
pp. 714-722 ◽  
Author(s):  
G. J. Hendricks ◽  
J. S. Sabnis ◽  
M. R. Feulner
Keyword(s):  
High Speed ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Rotating Stall ◽  
Two Dimensional ◽  
One Dimensional ◽  
High Pressure Ratio ◽  
Full Power ◽  
Dimensional Domain ◽  
Axial Flow Compressors

A nonlinear, two-dimensional, compressible dynamic model has been developed to study rotating stall/surge inception and development in high-speed, multistage, axial flow compressors. The flow dynamics are represented by the unsteady Euler equations, solved in each interblade row gap and inlet and exit ducts as two-dimensional domains, and in each blade passage as a one-dimensional domain. The resulting equations are solved on a computational grid. The boundary conditions between domains are represented by ideal turning coupled with empirical loss and deviation correlations. Results are presented comparing model simulations to instability inception data of an eleven stage, high-pressure-ratio compressor operating at both part and full power, and the results analyzed in the context of a linear modal analysis.

Centrifugal turbo chiller using water as refrigerant and lubricant

2020 ◽  
pp. 095440892093819
Author(s):  
Tadayoshi Shoyama ◽  
Bunki Kawano ◽  
Takeshi Ogata ◽  
Masaru Matsui ◽  
Masato Furukawa ◽  
...  
Keyword(s):  
Heat Pump ◽  
High Speed ◽  
Pressure Ratio ◽  
Control Valve ◽  
Cooling Capacity ◽  
Flow Control Valve ◽  
Pump System ◽  
Heat Pump System ◽  
High Pressure Ratio ◽  
Turbo Compressor

Water refrigerant heat pump system with a water vapour turbo compressor is developed. Water (R718) is an ideal refrigerant that is considered perfectly environment friendly. Although water refrigerant heat pump is studied extensively, the development of turbo compressors of high pressure ratio is still a technical challenge, in terms of both aerodynamics of the impellers and high-speed rotordynamics. In this study, the high-speed rotor is supported by journal bearings lubricated with water refrigerant. Additionally, other system components such as sprayed direct intercooler heat exchanger and anti-cavitation flow control valve have to be designed and developed with the new requirements as well. An experimental test rig of closed-loop heat pump is constructed which achieves the cooling capacity of 100 kW on the rated condition with COP = 5. The experimental results and the experienced challenges and their undertaken solutions are discussed in terms of the efficiency and the vibrations of the turbo compressor and the heat pump system.

Improvement of the Flow Range of Transonic Centrifugal Compressors With a Low-Solidity Cascade Diffuser

2000 ◽  
Author(s):  
Hiroshi Hayami
Keyword(s):  
Centrifugal Compressor ◽  
High Speed ◽  
Pressure Ratio ◽  
Input Power ◽  
Velocity Of Sound ◽  
Centrifugal Compressors ◽  
Vaned Diffuser ◽  
Transonic Centrifugal Compressor ◽  
Series Of Experiments ◽  
Low Solidity Cascade Diffuser

If the pressure ratio of a typical single-stage centrifugal compressor is larger than four, the velocity relative to the impeller and to the diffuser exceeds the velocity of sound. The flow range of transonic centrifugal compressors with a vaned diffuser is usually very narrow. Low-solidity cascade diffusers with solidity 0.69 have been successfully applied as a part of the diffuser system of a transonic centrifugal compressor. On the basis of this type of diffuser, a series of experiments to broaden the operating range are discussed focusing on the control of the geometry of impeller and/or diffuser; one was to reduce the inducer blade turning upstream of the throat, and the other was to reduce the inlet passage width of diffuser. The milder inducer blade camber realized the improvement in flow range by 1.5 times to the original one. Regarding the diffuser inlet passage width contraction, the flow range was not broadened so much owing to the change in impeller characteristics, but the input power was reduced and then the high speed efficiency was much improved.

The Use of the 3D-Viscous Codes in the Analysis of High Speed High Pressure Ratio Centrifugal Impellers

2000 ◽  
Author(s):  
Tarek Mekhail ◽  
Du Zhao Hui ◽  
Chen Han Ping ◽  
Willem Janson
Keyword(s):  
High Pressure ◽  
High Speed ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Internal Flow ◽  
Tip Clearance ◽  
Centrifugal Impeller ◽  
Field Calculation ◽  
High Pressure Ratio ◽  
Industrial Gas Turbine

Abstract The flow inside a centrifugal impeller has various complex three dimensional phenomena (flow separation, jet-wake structure, shock wave, etc.). In this study, the internal flow field calculation of Samsung, high pressure ratio, high speed, centrifugal impeller with splitter blades is obtained by commercially available CFX-Tascflow code with CFX-Turbogrid for grid generation. The results are compared to that obtained previously by Denton and Dawes codes. The impeller is used in the first stage centrifugal compressor of an industrial gas turbine. The CFX-Tascflow results showed some differences Mach number contours. Also, the calculations are performed for Krain’s backswept impeller and the results are compared to the experimental measurements. Simulation of tip clearance has been done and the results were in a good agreement with the previous experiments.

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.

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