Matrix Solution of Compressible Flow on S1 Surface Through a Turbomachine Blade Row With Splitter Vanes or Tandem Blades

1984 ◽  
Vol 106 (2) ◽  
pp. 449-454 ◽  
Author(s):  
Chung-Hua Wu ◽  
Baoguo Wang
Keyword(s):  
Compressible Flow ◽  
Present Method ◽  
Curvilinear Coordinates ◽  
Flow Ratio ◽  
Matrix Solution ◽  
Flow Angle ◽  
Trailing Edges ◽  
Blade Row ◽  
Outlet Flow ◽  
Turbomachine Blade

The basic aerothermodynamic equations of turbomachine flow expressed with respect to nonorthogonal curvilinear coordinates and corresponding nonorthogonal velocity components are used to solve the compressible flow S1 surface in a turbomachine blade row with splitter vanes or tandem blades. The equation of stream function is solved by matrix technique, and the mass flow ratio and outlet flow angle are determined by applying the Kutta-Joukowsky condition to the trailing edges of the main blade and the splitter vane. Typical examples are given to illustrate the effectiveness of the present method.

Multipoint optimization of an axial turbine cascade using a hybrid algorithm

2020 ◽  
pp. 1-15
Author(s):  
Arnaud Châtel ◽  
Tom Verstraete ◽  
Grégory Coussement
Keyword(s):  
Evolutionary Algorithm ◽  
Hybrid Algorithm ◽  
Search Space ◽  
Benchmark Problems ◽  
Flow Angle ◽  
Operating Range ◽  
Trailing Edges ◽  
Gradient Based ◽  
Outlet Flow ◽  
Multipoint Optimization

Abstract This paper presents a multipoint optimization of the LS89 cascade. The objective of the optimization consists in minimizing the entropy losses generated inside the cascade over a predefined operating range. Two aerodynamic constraints are imposed in order to conserve the same performance as the original cascade. The first constraint is established on the outlet flow angle in order to achieve at least the same flow turning as the LS89. The second constraint limits the mass-flow passing through the cascade. The optimization is performed using a hybrid algorithm which combines a classical evolutionary algorithm with a gradient-based method. The hybridization between both methods is based on the Lamarckian approach which consists in incorporating the gradient method inside the loop of the evolutionary algorithm. In this methodology, the evolutionary method allows to globally explore the design space while the gradient-based method locally improves certain designs located in promising regions of the search space. First, the better performance of the hybrid method compared to the performance of an evolutionary algorithm is demonstrated on benchmark problems. Then, the methodology is applied on the LS89 application. The optimization allows to find a new profile which reduces the entropy losses over the entire operating range by at least 9.5 %. Finally, the comparison of the flows computed in the baseline and in the optimized cascades demonstrates that the reduction of the losses is due to a decrease of the entropy generated downstream the trailing edges and within the passages between the optimized blades.

Experimental assessment of the rotor outlet flow in a twin-entry radial turbine by means of Laser Doppler Anemometry

2021 ◽  
pp. 146808742110344
Author(s):  
José Galindo ◽  
Andrés Omar Tiseira ◽  
Luis Miguel García-Cuevas ◽  
Nicolás Medina
Keyword(s):  
Laser Doppler Anemometry ◽  
Effective Area ◽  
Laser Doppler ◽  
Flow Ratio ◽  
One Dimensional ◽  
Radial Turbine ◽  
Single Entry ◽  
Outlet Flow ◽  
Dimensional Models ◽  
Number Of Particles

The current paper presents the validation of some hypotheses used for developing a one-dimensional twin-entry turbine model with experimental measurements. A Laser Doppler Anemometry (LDA) technique has been used for measuring the axial Mach number and for counting the number of particles downstream of the rotor outlet. These measurements have been done for different mass flow ratio (MFR) and reduced turbocharger speed conditions. The flow coming from each turbine entry does not fully mix with the other within the rotor since, downstream of the rotor, they can still be differentiated. Thus, the hypothesis of studying twin-entry turbines as two separated single-entry turbines in one-dimensional models is corroborated. Moreover, the rotor outlet area corresponding to each flow branch has linear trends with the MFR value. Therefore, the rotor outlet effective area used for one-dimensional models should vary linearly with the MFR value.

Nonstaggered APPLE Algorithm for Incompressible Viscous Flow in Curvilinear Coordinates

2002 ◽  
Vol 124 (5) ◽  
pp. 812-819 ◽  
Author(s):  
S. L. Lee ◽  
Y. F. Chen
Keyword(s):  
Viscous Flow ◽  
Present Method ◽  
Heat Conduction Equation ◽  
Continuity Equation ◽  
Interpolation Method ◽  
Curvilinear Coordinates ◽  
Grid System ◽  
Weighted Interpolation ◽  
Incompressible Viscous Flow ◽  
Curvilinear Grid

The NAPPLE algorithm for incompressible viscous flow on Cartesian grid system is extended to nonorthogonal curvilinear grid system in this paper. A pressure-linked equation is obtained by substituting the discretized momentum equations into the discretized continuity equation. Instead of employing a velocity interpolation such as pressure-weighted interpolation method (PWIM), a particular approximation is adopted to circumvent the checkerboard error such that the solution does not depend on the under-relaxation factor. This is a distinctive feature of the present method. Furthermore, the pressure is directly solved from the pressure-linked equation without recourse to a pressure-correction equation. In the use of the NAPPLE algorithm, solving the pressure-linked equation is as simple as solving a heat conduction equation. Through two well-documented examples, performance of the NAPPLE algorithm is validated for both buoyancy-driven and pressure-driven flows.

Control of Shroud Leakage Loss by Reducing Circumferential Mixing

2008 ◽  
Vol 130 (2) ◽  
Author(s):  
Budimir Rosic ◽  
John D. Denton
Keyword(s):  
Significant Proportion ◽  
Tangential Velocity ◽  
Secondary Flows ◽  
Main Stream ◽  
Leakage Flow ◽  
Flow Angle ◽  
Blade Row ◽  
Stator Blade ◽  
The Difference ◽  
Mixing Losses

Shroud leakage flow undergoes little change in the tangential velocity as it passes over the shroud. Mixing due to the difference in tangential velocity between the main stream flow and the leakage flow creates a significant proportion of the total loss associated with shroud leakage flow. The unturned leakage flow also causes negative incidence and intensifies the secondary flows in the downstream blade row. This paper describes the experimental results of a concept to turn the rotor shroud leakage flow in the direction of the main blade passage flow in order to reduce the aerodynamic mixing losses. A three-stage air model turbine with low aspect ratio blading was used in this study. A series of different stationary turning vane geometries placed into the rotor shroud exit cavity downstream of each rotor blade row was tested. A significant improvement in flow angle and loss in the downstream stator blade rows was measured together with an increase in turbine brake efficiency of 0.4 %.

scholarly journals Effects of Transient Heat Transfer on Compressor Stability

2018 ◽  
Vol 140 (12) ◽  
Author(s):  
A. Kiss ◽  
Z. Spakovszky
Keyword(s):  
Heat Transfer ◽  
Pressure Ratio ◽  
Strong Dependence ◽  
Transient Heat Transfer ◽  
Compressor Blades ◽  
Flow Angle ◽  
Stall Margin ◽  
Point Reduction ◽  
Angle Deviation ◽  
Blade Row

The effects of heat transfer between the compressor structure and the primary gas path flow on compressor stability are investigated during hot engine re-acceleration transients. A mean line analysis of an advanced, high-pressure ratio compressor is extended to include the effects of heat transfer on both stage matching and blade row flow angle deviation. A lumped capacitance model is used to compute the heat transfer of the compressor blades, hub, and casing to the primary gas path. The inputs to the compressor model with heat transfer are based on a combination of full engine data, compressor test rig measurements, and detailed heat transfer computations. Nonadiabatic transient calculations show a 8.0 point reduction in stall margin from the adiabatic case, with heat transfer predominantly altering the transient stall line. 3.4 points of the total stall margin reduction are attributed to the effect of heat transfer on blade row deviation, with the remainder attributed to stage rematching. Heat transfer increases loading in the front stages and destabilizes the front block. Sensitivity studies show a strong dependence of stall margin to heat transfer magnitude and flow angle deviation at low speed, due to the effects of compressibility. Computations for the same transient using current cycle models with bulk heat transfer effects only capture 1.2 points of the 8.0 point stall margin reduction. Based on this new capability, opportunities exist early in the design process to address potential stability issues due to transient heat transfer.

Forces and Surface Pressure on a Blade Moving in Front of the Admission Region

2010 ◽  
Vol 132 (12) ◽  
Author(s):  
Soo-Yong Cho ◽  
Chong-Hyun Cho ◽  
Kook-Young Ahn ◽  
Young-Cheol Kim
Keyword(s):  
Cross Section ◽  
Performance Test ◽  
Nozzle Flow ◽  
Flow Angle ◽  
Linear Cascade ◽  
Blade Row ◽  
Partial Admission ◽  
Additional Losses ◽  
Rotational Direction ◽  
Admission Region

The partial admission technique is widely used to control the output power of turbines. In some cases, it has more merits than full admission. However, additional losses, such as expansion, mixing, or pumping, are generated in partial admission as compared with full admission. Thus, an experiment was conducted in a linear cascade apparatus having a partial admission region in order to investigate the effect of partial admission on a blade row. The admission region was formed by a spouting nozzle installed at the inlet of the linear cascade apparatus. Its cross section was rectangular and its size is 200×200 mm2. The tested blade was axial-type and its chord was 200 mm. Nineteen identical blades were applied to the linear cascade for the partial admission experiment. The blades moved along the rotational direction in front of the admission region, and then operating forces and surface pressures on the blades were measured at the steady state. The experiment was conducted at a Reynolds number of 3×105 based on the chord. The nozzle flow angle was set to 65 deg with a solidity of 1.38 for performance test at the design point. In addition, another two different solidities of 1.25 and 1.67 were applied. From the experimental results, when the solidity was decreased, the maximum rotational force increased but the maximum axial force decreased.

Cooling-Air Injection Into Secondary Flow and Loss Fields Within a Linear Turbine Cascade

1991 ◽  
Vol 113 (3) ◽  
pp. 375-383 ◽  
Author(s):  
A. Yamamoto ◽  
Y. Kondo ◽  
R. Murao
Keyword(s):  
Leading Edge ◽  
Secondary Flows ◽  
Air Injection ◽  
Internal Flows ◽  
Flow Angle ◽  
Secondary Air Injection ◽  
Outlet Flow ◽  
Overall Performance ◽  
Secondary Air ◽  
Cooling Air

In order to understand overall performance and internal flows of air-cooled turbine blade rows, flows in a model linear cascade were surveyed with secondary air injection from various locations of the blade surfaces. The secondary air interacted with the cascade passage vortices and changed the loss distribution significantly. The cascade overall loss decreased when the air was injected along the mainstream and increased when the air was injected against the mainstream from some locations of the blade leading edge. Effects on overall kinetic energy of the secondary flows and on the cascade outlet flow angle were also discussed in this paper.

Evaluation of Unsteady CFD Methods by Their Application to a Transonic Propfan Stage

2001 ◽  
Author(s):  
S. Schmitt ◽  
F. Eulitz ◽  
L. Wallscheid ◽  
A. Arnone ◽  
M. Marconcini
Keyword(s):  
Numerical Methods ◽  
Navier Stokes ◽  
Test Case ◽  
Flow Angle ◽  
Laser Measurements ◽  
General Flow ◽  
Blade Row ◽  
Flow Features ◽  
Blade Row Interaction ◽  
Good Agreement

The accuracy in predicting the unsteady aerodynamic blade-row-interaction of two state-of-the-art Navier-Stokes codes is evaluated within the current paper. The general flow features of the test case — a transonic research propfan stage — are described in brief as far as necessary to understand the detailed comparisons. The calculated unsteady velocity and flow angle distributions at various axial planes of the stage are compared to data from unsteady laser measurements. The general flow features of the propfan are very well reproduced by the numerical methods and a good agreement is also obtained in comparison to the measured data. One important outcome of the comparison is the good agreement of both numerical methods with the unsteady fluctuations measured in the experiment.

The Relation Between Solidity and the Other Shape Factors for Complete Optimum Meridian Profile of Impeller With Guidevane

2007 ◽  
Author(s):  
Takuji Tsugawa
Keyword(s):  
Shape Factor ◽  
Large Diameter ◽  
Axial Direction ◽  
The Other ◽  
Shape Factors ◽  
Flow Angle ◽  
Diffusion Factor ◽  
Blade Number ◽  
Outlet Flow ◽  
Angle Blade

In the previous paper, the solidity is independent shape factor of the optimum meridian profile by diffusion factor. But, the solidity is often calculated by the other shape factors, for example, the inlet and outlet flow angle, blade length, blade number and the co-ordinates of impeller meridian profile. So, in this paper, the solidity is treated as dependent shape factor and is calculated by the impeller meridian co-ordinates and flow angle. In the previous paper, the impeller meridian inlet is axial direction. In this paper, the inlet mixed flow angle of impeller inlet is one of additional shape factor. As the result, the impeller with guidevane complete meridian profile is calculated for the large diameter of guidevane outlet and the detailed meridian profile of impeller inlet.

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