Transonic and Low Supersonic Flow Losses of Two Steam Turbine Blades at Large Incidences

2004 ◽  
Vol 126 (6) ◽  
pp. 966-975 ◽  
Author(s):  
S.-M. Li ◽  
T.-L. Chu ◽  
Y.-S. Yoo ◽  
W. F. Ng
Keyword(s):  
Experimental Data ◽  
Supersonic Flow ◽  
Turbine Blades ◽  
Supersonic Flows ◽  
Relative Increase ◽  
Pressure Probe ◽  
Flow Losses ◽  
New Correlation ◽  
Turbine Cascades ◽  
Mach Numbers

A linear cascade experiment was conducted to investigate transonic and low supersonic flow losses of two nozzle blades for the steam turbines. In the experiment, flow incidences were changed from −34° to 35° and exit Mach numbers were varied from 0.60 to 1.15. Tests were conducted at Reynolds numbers between 7.4×105 and 1.6×106. Flow visualization techniques, such as shadowgraph, Schlieren, and surface color oil were used to document the flows. Measurements were made by using downstream traverses with Pitot probe, upstream total pressure probe, and sidewall static pressure taps. The losses were found to be rather constant at subsonic flows. At transonic and low supersonic flows, the losses increased steeply. The maximum relative increase of the losses was near 700% when the Mach numbers increased from 0.6 to 1.15. However, the maximum relative increase of the losses was only about 100% due to very large variation of incidences. It is important to note that the effect of Mach numbers on losses was much greater than that due to the very large incidences for the transonic and low supersonic flows. A frequently used loss correlation in the literature is found not suitable to predict the losses of the tested blades for the transonic and low supersonic flows. From the current experimental data and some data in the literature, a new correlation for the shock related losses is proposed for transonic and low supersonic flows of turbine cascades. Comparison is made among the existing correlation and the new correlation, as well as the data of the current two cascades and other three turbine cascades in the literature. Improved agreement with the experimental data of the five cascades is obtained by using the new correlation as compared with the prediction by using the frequently used loss correlation in the literature.

Transonic Flow Losses of Two Steam Turbine Blades at Large Incidences

2002 ◽  
Author(s):  
S.-M. Li ◽  
T.-L. Chu ◽  
Y.-S. Yoo ◽  
W. F. Ng
Keyword(s):  
Shock Wave ◽  
Transonic Flow ◽  
Turbine Blades ◽  
Supersonic Flows ◽  
Pressure Probe ◽  
Flow Measurements ◽  
Subsonic Flows ◽  
Flow Losses ◽  
Mach Numbers ◽  
Loss Level

A linear cascade experiment has been conducted to investigate transonic flow losses of two nozzle blade profiles for high-pressure steam turbines. Experimental incidences were changed from −34° to 35°, while exit Mach numbers were varied from 0.60 to 1.15. Reynolds numbers were tested between 7.4×105 and 1.6×106. Flow visualization techniques, such as shadowgraph, Schlieren, and surface color oil were used to document the flow. Measurements were made using downstream traverses with Pitot probe, upstream total pressure probe, and sidewall static pressure taps. Losses were found to be rather constant at subsonic flows. At transonic and low supersonic flows, the loss increased steeply, as high as 700% more when compared with the loss level at subsonic flows. For all the tested Mach numbers, the maximum relative increase of loss level was only about 100% due to the very large variation of incidences. Thus the effect of Mach numbers on losses is much greater than that of the very large incidences for the transonic and low supersonic flows. A frequently used shock wave loss correlation was found not suitable to predict the losses of the tested blades. Using the current experimental data, a modified loss correlation to account for the shock wave effects has been proposed for the transonic and low supersonic flows. Comparison was made between the modified loss correlation and the data of three other turbine cascades in the literature. Large improvement was obtained with the modified loss correlation as compared with the existing correlation.

scholarly journals The Chordwise Lean of Highly Loaded Turbine Blades

1996 ◽  
Author(s):  
Zhongqi Wang ◽  
Wanjin Han
Keyword(s):  
Experimental Data ◽  
Experimental Investigation ◽  
Saddle Point ◽  
Turbine Blades ◽  
Acute Angle ◽  
Secondary Vortex ◽  
Pressure Gradients ◽  
Turbine Cascades ◽  
Highly Loaded

An experimental investigation was carried out on the effect of blade chordwise lean on the losses in highly loaded rectangular turbine cascades. Detail measurements include 10 traverses from the upstream to the downstream of the cascades with five-hole spherical probes. Compared with the experimental data of the conventional straight and pitchwise lean blades under the same conditions, it is shown that the effect of chordwise lean on the development of the cascade losses is similar to that of pitchwise lean. However, the chordwise lean produces smaller streamwise adverse pressure gradients near both endwalls and a smaller spanwise negative one starting from the acute angle side in the first part of the passages in chordwise lean cascade, thereby the saddle point separations and intensities of the passage vortices are weakened and the secondary vortex losses are cut down notably.

scholarly journals Investigation of gasdynamic parameters of a supersonic flow near a body at various Mach numbers

2020 ◽  
Vol 1697 ◽  
pp. 012240
Author(s):  
A V Shevchenko ◽  
A S Yuriev ◽  
S A Poniaev ◽  
T A Zhitnikov ◽  
E B Panfilov ◽  
...  
Keyword(s):  
Supersonic Flow ◽  
Mach Numbers

Validating Numerical CFD Simulations With Experimental Data for Turbulence Phenomena in Axial Flow Gas Turbine Diffusers

2009 ◽  
Author(s):  
Farrokh Zarifi-Rad ◽  
Hamid Vajihollahi ◽  
James O’Brien
Keyword(s):  
Experimental Data ◽  
Gas Turbine ◽  
Turbine Blades ◽  
Axial Flow ◽  
Experimental Results ◽  
Scale Model ◽  
Data Set ◽  
Pressure Profiles ◽  
Scale Models ◽  
Inlet Conditions

Scale models give engineers an excellent understanding of the aerodynamic behavior behind their design; nevertheless, scale models are time consuming and expensive. Therefore computer simulations such as Computational Fluid Dynamics (CFD) are an excellent alternative to scale models. One must ask the question, how close are the CFD results to the actual fluid behavior of the scale model? In order to answer this question the engineering team investigated the performance of a large industrial Gas Turbine (GT) exhaust diffuser scale model with performance predicted by commercially available CFD software. The experimental results were obtained from a 1:12 scale model of a GT exhaust diffuser with a fixed row of blades to simulate the swirl generated by the last row of turbine blades five blade configurations. This work is to validate the effect of the turbulent inlet conditions on an axial diffuser, both on the experimental front and on the numerical analysis approach. The object of this work is to bring forward a better understanding of velocity and static pressure profiles along the gas turbine diffusers and to provide an accurate experimental data set to validate the CFD prediction. For the CFD aspect, ANSYS CFX software was chosen as the solver. Two different types of mesh (hexagonal and tetrahedral) will be compared to the experimental results. It is understood that hexagonal (HEX) meshes are more time consuming and more computationally demanding, they are less prone to mesh sensitivity and have the tendancy to converge at a faster rate than the tetrahedral (TET) mesh. It was found that the HEX mesh was able to generate more consistent results and had less error than TET mesh.

A correlation to predict the performance characteristics of centrifugal pumps handling slurries

1997 ◽  
Vol 211 (2) ◽  
pp. 147-157 ◽  
Author(s):  
K A Kazim ◽  
B Maiti ◽  
P Chand
Keyword(s):  
Experimental Data ◽  
Particle Size ◽  
Suspended Solids ◽  
Reduction Factor ◽  
Performance Characteristics ◽  
Centrifugal Pumps ◽  
Pump Performance ◽  
New Correlation ◽  
Effect Of Particle Size ◽  
The Individual

Centrifugal pumps are being used increasingly for transportation of slurries through pipelines. To design a slurry handling system it is essential to have a knowledge of the effects of suspended solids on the pump performance. A new correlation to predict the head reduction factor for centrifugal pumps handling solids has been developed. This correlation takes into account the individual effect of particle size, particle size distribution, specific gravity and concentration of solids on the centrifugal pump performance characteristics. The range of validity of the correlation has been verified by experiment and by using experimental data available from the literature. The present correlation shows better agreement with the experimental data than existing correlations.

Analysis of Blunt Trailing Edge Airfoils

2003 ◽  
Author(s):  
K. J. Standish ◽  
C. P. van Dam
Keyword(s):  
Experimental Data ◽  
Wind Turbine ◽  
Turbine Blades ◽  
Navier Stokes ◽  
Computational Techniques ◽  
Wind Turbine Blades ◽  
Trailing Edge ◽  
Blunt Trailing Edge ◽  
Large Wind ◽  
Structural Volume

The adoption of blunt trailing edge airfoils for the inner regions of large wind turbine blades has been proposed. Blunt trailing edge airfoils would not only provide increased structural volume, but have also been found to improve the lift characteristics of airfoils and therefore allow for section shapes with a greater maximum thickness. Limited experimental data makes it difficult for wind turbine designers to consider and conduct tradeoff studies using these section shapes. This lack of experimental data precipitated the present analysis of blunt trailing edge airfoils using computational fluid dynamics. Several computational techniques are applied including a viscous/inviscid interaction method and several Reynolds-averaged Navier-Stokes methods.

Computational study of unsteady gas flow in the combustion chamber of a ramjet engine with heat transfer

2021 ◽  
pp. 50-61
Author(s):  
Надежда Петровна Скибина
Keyword(s):  
Heat Transfer ◽  
Numerical Simulation ◽  
Supersonic Flow ◽  
Combustion Chamber ◽  
Gas Flow ◽  
Stokes Equations ◽  
Heat Conducting ◽  
Measuring Device ◽  
Ramjet Engine ◽  
Mach Numbers

Проведено численное исследование нестационарного турбулентного сверхзвукового течения в камере сгорания прямоточного воздушно-реактивного двигателя. Описана методика экспериментального измерения температуры на стенке осесимметричного канала в камере сгорания двигателя. Математическое моделирование обтекания исследуемой модели двигателя проводилось для скоростей набегающего потока M = 5 ... 7. Начальные и граничные условия задачи соответствовали реальному аэродинамическому эксперименту. Проанализированы результаты численного расчета. Рассмотрено изменение распределения температуры вдоль стенки канала с течением времени. Проведена оценка согласованности полученных экспериментальных данных с результатами математического моделирования. Purpose. The aim of this study is a numerical simulation of unsteady supersonic gas flow in a working path of ramjet engine under conditions identical to aerodynamic tests. Free stream velocity corresponding to Mach numbers M=5 ... 7 are considered. Methodology. Presented study addresses the methods of physical and numerical simulation. The probing device for thermometric that allows to recording the temperature values along the wall of internal duct was proposed. To describe the motion of a viscous heat-conducting gas the unsteady Reynolds averaged Navier - Stokes equations are considered. The flow turbulence is accounted by the modified SST model. The problem was solved in ANSYS Fluent using finite-volume method. The initial and boundary conditions for unsteady calculation are set according to conditions of real aerodynamic tests. The coupled heat transfer for supersonic flow and elements of ramjet engine model are realized by setting of thermophysical properties of materials. The reliability testing of numerical simulation has been made to compare the results of calculations and the data of thermometric experimental tests. Findings. Numerical simulation of aerodynamic tests for ramjet engine was carried out. The agreement between the results of numerical calculations and experimental measurements for the velocity in the channel under consideration was obtained; the error was shown to be 2%. The temperature values were obtained in the area of contact of the supersonic flow with the surface of the measuring device for the external incident flow velocities for Mach numbers M = 5 ... 7. The process of heating the material in the channel that simulated the section of the engine combustion chamber was analyzed. The temperature distribution was studied depending on the position of the material layer under consideration relative to the contact zone with the flow. Value. In the course of the work, the fields of flow around the model of a ramjet engine were obtained, including the region of supersonic flow in the inner part of axisymmetric channel. The analysis of the temperature fields showed that to improve the quality of the results, it is necessary to take into account the depth of the calorimetric sensor. The obtained results will be used to estimate the time of interaction of the supersonic flow with the fuel surface required to reach the combustion temperature.

Experimental study of a pitching and plunging wing

2018 ◽  
Vol 90 (7) ◽  
pp. 1136-1144 ◽  
Author(s):  
Dimitris Gkiolas ◽  
Demetri Yiasemides ◽  
Demetri Mathioulakis
Keyword(s):  
Experimental Data ◽  
Flow Visualization ◽  
Wind Turbine ◽  
Flow Separation ◽  
Flow Behavior ◽  
Lift Coefficient ◽  
Turbine Blades ◽  
Content Type ◽  
Separation Line ◽  
Wing Chord

Purpose The complex flow behavior over an oscillating aerodynamic body, e.g. a helicopter rotor blade, a rotating wind turbine blade or the wing of a maneuvering airplane involves combinations of pitching and plunging motions. As the parameters of the problem (Re, St and phase difference between these two motions) vary, a quasi-steady analysis fails to provide realistic results for the aerodynamic response of the moving body, whereas this study aims to provide reliable experimental data. Design/methodology/approach In the present study, a pitching and plunging mechanism was designed and built in a subsonic closed-circuit wind tunnel as well as a rectangular aluminum wing of a 2:1 aspect-ratio with a NACA64-418 airfoil, used in wind turbine blades. To measure the pressure distribution along the wing chord, a number of fast responding transducers were embedded into the mid span wing surface. Simultaneous pressure measurements were conducted along the wing chord for the Reynolds number of 0.85 × 106 for both steady and unsteady cases (pitching and plunging). A flow visualization technique was used to detect the flow separation line under steady conditions. Findings Elevated pressure fluctuations coincide with the flow separation line having been detected through surface flow visualization and flattened pressure distributions appear downstream of the flow separation line. Closed hysteresis loops of the lift coefficient versus angle of attack were measured for combined pitching and plunging motions. Practical implications The experimental data can be used for improvement of unsteady fluid mechanics problem solvers. Originality/value In the present study, a new installation was built allowing the aerodynamic study of oscillating wings performing pitching and plunging motions with prescribed frequencies and phase lags between the two motions. The experimental data can be used for improvement of computational fluid dynamics codes in case that the examined aerodynamic body is oscillating.

Experimental Validation of Forced Response Methods in a Multi-Stage Axial Turbine

2018 ◽  
Author(s):  
Thomas Hauptmann ◽  
Christopher E. Meinzer ◽  
Joerg R. Seume
Keyword(s):  
Experimental Data ◽  
Vibration Amplitude ◽  
Turbine Blades ◽  
Service Condition ◽  
Sensitivity Study ◽  
Forced Response ◽  
Tip Clearance ◽  
Computational Time ◽  
Reference Case ◽  
Axial Turbine

Depending on the in service condition of jet engines, turbine blades may have to be replaced, refurbished, or repaired in the course of an engine overhaul. Thus, significant changes of the turbine blade geometry can be introduced due to regeneration and overhaul processes. Such geometric variances can affect the aerodynamic and aeroelastic behavior of turbine blades. One goal in the development of the regeneration process is to estimate the aerodynamic excitation of turbine blades depending on these geometric variances caused during the regeneration. Therefore, this study presents an experimentally validated comparison of two methods for the prediction of forced response in a multistage axial turbine. Two unidirectional fluid structure interaction (FSI) methods, a time-linearized and a time-accurate with a subsequent linear harmonic analysis, are employed and the results validated against experimental data. The results show that the vibration amplitude of the time-linearized method is in good agreement with the experimental data and, also requires lower computational time than the time-accurate FSI. Based on this result, the time-linearized method is used to perform a sensitivity study of the tip clearance size of the last rotor blade row of the five stage axial turbine. The results show that an increasing tip clearances size causes an up to 1.35 higher vibration amplitude compared to the reference case, due to increased forcing and decreased damping work.

Loss Generation in Transonic Turbine Blading

2017 ◽  
Author(s):  
Penghao Duan ◽  
Choon S. Tan ◽  
Andrew Scribner ◽  
Anthony Malandra
Keyword(s):  
Experimental Data ◽  
Boundary Layer ◽  
Mach Number ◽  
High Speed ◽  
Turbine Blades ◽  
Surface Boundary ◽  
Loss Model ◽  
Exit Mach Number ◽  
Loss Characteristic ◽  
Shock Loss

The measured loss characteristic in a high-speed cascade tunnel of two turbine blades of different designs showed distinctly different trend with exit Mach number ranging from 0.8 to 1.4. Assessments using steady RANS computation of the flow in the two turbine blades, complemented with control volume analyses and loss modelling, elucidate why the measured loss characteristic looks the way it is. The loss model categorizes the total loss in terms of boundary layer loss, trailing edge loss and shock loss; it yields results in good agreement with the experimental data as well as steady RANS computed results. Thus RANS is an adequate tool for determining the loss variations with exit isentropic Mach number and the loss model serves as an effective tool to interpret both the computational and experimental data. The measured loss plateau in Blade 1 for exit Mach number of 1 to 1.4 is due to a balance between a decrease of blade surface boundary layer loss and an increase in the attendant shock loss with Mach number; this plateau is absent in Blade 2 due to a greater rate in shock loss increase than the corresponding decrease in boundary layer loss. For exit Mach number from 0.85 to 1, the higher loss associated with shock system in Blade 1 is due to the larger divergent angle downstream of the throat than that in Blade 2. However when exit Mach number is between 1.00 and 1.30, Blade 2 has higher shock loss. For exit Mach number above around 1.4, the shock loss for the two blades is similar as the flow downstream of the throat is completely supersonic. In the transonic to supersonic flow regime, the turbine design can be tailored to yield a shock pattern the loss of which can be mitigated in near equal amount of that from the boundary layer with increasing exit Mach number, hence yielding a loss plateau in transonic-supersonic regime.

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