Experimental Investigation of Secondary Flow and Mixing Downstream of Straight Turbine Cascades

1988 ◽  
Vol 110 (4) ◽  
pp. 497-503 ◽  
Author(s):  
A. Mobarak ◽  
M. G. Khalafallah ◽  
A. M. Osman ◽  
H. A. Heikal
Keyword(s):  
Boundary Layer ◽  
Experimental Investigation ◽  
Mach Number ◽  
Aspect Ratio ◽  
Energy Loss ◽  
Total Energy ◽  
Secondary Flow ◽  
Boundary Layer Thickness ◽  
Turbine Cascades ◽  
Exit Mach Number

The purpose of this paper is to investigate the flow field downstream of turbine cascades of low aspect ratio, often used in vehicles and small turbomachines. Experimental investigation was carried out to study the flow downstream of three sets of turbine cascades having the same blade turning angle of about 83 deg but different profiles. The total energy losses were measured at several planes downstream of the cascade of blades in order to determine the changes in gross secondary flow loss coefficient and the growth of the mixing loss with distance downstream. Influence of inlet boundary layer thickness, aspect ratio, and exit Mach number on the nature of the flow at the exit plane of the cascade and total energy loss were studied. The tests were performed with four values of aspect ratio: 1.16, 0.8, 0.5, and 0.25. Some new correlations were deduced that predict energy loss coefficients as a function of distance downstream, aspect ratio, and exit Mach number as well as the upstream boundary layer thickness. The test results compare well with other published correlations.

scholarly journals Investigation of the effects of main geometric parameters and flow characteristics on secondary flow losses in a turbine cascade

2021 ◽  
Vol 2131 (3) ◽  
pp. 032081
Author(s):  
M Mesbah ◽  
V G Gribin ◽  
K Souri
Keyword(s):  
Mach Number ◽  
Aspect Ratio ◽  
Total Energy ◽  
Flow Characteristics ◽  
Numerical Results ◽  
Energy Losses ◽  
Inlet Flow ◽  
Flow Angle ◽  
Flow Losses ◽  
Exit Mach Number

Abstract This paper presents numerical simulation results of a three-dimensional (3D) transitional flow in a stator cascade of an axial turbine. The influences of the main geometric parameters and flow characteristics including, the blade aspect ratio, pitch-to-chord ratio, inlet flow angle, and exit Mach number, on secondary flows development and end-wall losses, were studied. The numerical results were validated by the results of experiments conducted in the laboratory of the steam and gas turbine faculty of the Moscow Power Engineering Institute. The maximum difference between computed and experimental results was 2.4 %. The total energy losses decrease by 20 % when the exit Mach number changes from 0.38 to 0.8. Numerical results indicated that the blade aspect ratio had the most effect on secondary flow losses. The total energy losses increase by 46.6 % when the aspect ratio decreases from 1 to 0.25. The total loss of energy by 13.2 % decreases by increasing the inlet flow angle from 60 degrees to 90 degrees. Then by increasing the inlet flow angle from 90 to 110 degrees, the total loss rises by 3.6%. As the pitch-to-chord ratio increases from 0.7 to 0.75, the total energy losses are reduced by 12.2 %. Then by increasing the pitch-to-chord ratio from 0.75 to 0.8, the total energy losses increase by 6 %. As with experimental data, the numerical results showed that the optimal inlet flow angle and relative pitch for the cascade are 90 degrees and 0.75, respectively.

Some Boundary Layer Measurements on a Slender Wing at Supersonic Speeds

1963 ◽  
Vol 67 (629) ◽  
pp. 291-295
Author(s):  
R. T. Griffiths
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Aspect Ratio ◽  
Flat Plate ◽  
Layer Thickness ◽  
Incompressible Flow ◽  
Boundary Layer Thickness ◽  
Static Pressure ◽  
Leading Edge ◽  
Slender Wing

SummaryBoundary layer measurements have been made at four positions on a slender gothic wing of aspect ratio 0·75. Test's were made over a range of incidence at M=1·42 and 1·82. With transition fixed by roughness near the leading edge the boundary layer thickness varied little with small positive or negative incidence but was reduced at larger incidences, this being most marked at positive incidence for positions nearest the leading edge due to the influence of the wing vortex. With the exception of positions in the vicinity of the vortex, a good estimate of the boundary layer thickness was given by the theory for incompressible flow over a flat plate and an excellent estimate of the variation of local static pressure and Mach number with incidence was given by not-so-slender wing theory.

scholarly journals Study of Various Factors Affecting Secondary Loss Vortices Downstream a Straight Turbine Cascade

1989 ◽  
Author(s):  
A. Mobarak ◽  
M. G. Khalafallah ◽  
H. A. Heikal ◽  
A. M. Osman
Keyword(s):  
Boundary Layer ◽  
Experimental Investigation ◽  
Boundary Layer Thickness ◽  
Turbine Cascade ◽  
Turning Angle ◽  
Factors Affecting ◽  
Secondary Loss ◽  
Downstream Distance ◽  
Exit Mach Number ◽  
Secondary Vortices

An experimental investigation is carried out to study the effect of inlet boundary layer thickness, aspect ratio and exit Mach number on the energy loss and on the behaviour of secondary vortices downstream of a straight turbine cascade of blades having 82° turning angle. Some correlations were deduced which predict, the cascade efficiency and the gross secondary loss coefficient as function of the downstream distance.

scholarly journals The sensitivity of turbine cascade endwall loss to inlet boundary layer thickness

2019 ◽  
Vol 3 ◽  
pp. OEYMDE ◽  
Author(s):  
John Coull ◽  
Christopher Clark ◽  
Raul Vazquez
Keyword(s):  
Boundary Layer ◽  
Secondary Flow ◽  
Layer Thickness ◽  
Boundary Layer Thickness ◽  
Flow Theory ◽  
Sensitivity Assessment ◽  
Blade Row ◽  
Turbine Cascades ◽  
Inlet Conditions ◽  
Is Design

The development of hub and casing boundary layers through a turbomachine is difficult to predict, giving rise to uncertainty in the boundary conditions experienced by each blade row. Previous studies in turbine cascades disagree on the sensitivity of endwall loss to such inlet conditions. This paper explores the problem computationally, by examining a large number of turbine cascades and varying the inlet boundary layer thickness. It is demonstrated that the sensitivity of endwall loss to inlet conditions is design dependent, and determined by the component of endwall loss associated with the secondary flow. This Secondary-Flow-Induced loss is characterised by a vorticity factor based on classical secondary flow theory. Designs that produce high levels of secondary vorticity tend to generate more loss and are more sensitive to inlet conditions. This sensitivity is largely driven by the dissipation of Secondary Kinetic Energy (SKE): thickening the inlet boundary layer causes the secondary vorticity at the cascade exit to be more dispersed within the passage, resulting in larger secondary flow structures with higher SKE. The effects are captured using a simple streamfunction model based on classical secondary flow theory, which has potential for preliminary design and sensitivity assessment.

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.

Effects of Shock Wave Development on Secondary Flow Behavior in Linear Turbine Cascade at Transonic Condition

2020 ◽  
Author(s):  
Hoshio Tsujita ◽  
Masanao Kaneko
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Secondary Flow ◽  
Turbine Blade ◽  
Gas Turbines ◽  
Flow Behavior ◽  
Turbine Cascade ◽  
Blade Loading ◽  
Passage Vortex ◽  
Exit Mach Number

Abstract Gas turbines widely applied to power generation and aerospace propulsion systems are continuously enhanced in efficiency for the reduction of environmental load. The energy recovery efficiency from working fluid in a turbine component constituting gas turbines can be enhanced by the increase of turbine blade loading. However, the increase of turbine blade loading inevitably intensifies the secondary flows, and consequently increases the associated loss generation. The development of the passage vortex is strongly influenced by the pitchwise pressure gradient on the endwall in the cascade passage. In addition, a practical high pressure turbine stage is generally driven under transonic flow conditions where the shock wave strongly influences the pressure distribution on the endwall. Therefore, it becomes very important to clarify the effects of the shock wave formation on the secondary flow behavior in order to increase the turbine blade loading without the deterioration of efficiency. In this study, the two-dimensional and the three-dimensional transonic flows in the HS1A linear turbine cascade at the design incidence angle were analyzed numerically by using the commercial CFD code with the assumption of steady compressible flow. The isentropic exit Mach number was varied from the subsonic to the supersonic conditions in order to examine the effects of development of shock wave caused by the increase of exit Mach number on the secondary flow behavior. The increase of exit Mach number induced the shock across the passage and increased its obliqueness. The increase of obliqueness reduced the cross flow on the endwall by moving the local minimum point of static pressure along the suction surface toward the trailing edge. As a consequence, the increase of exit Mach number attenuated the passage vortex.

A Review of Cascade Data on Secondary Losses in Turbines

1970 ◽  
Vol 12 (1) ◽  
pp. 48-59 ◽  
Author(s):  
J. Dunham
Keyword(s):  
Boundary Layer ◽  
Layer Thickness ◽  
Boundary Layers ◽  
Boundary Layer Thickness ◽  
Blade Loading ◽  
Axial Turbine ◽  
Wall Boundary Layer ◽  
Wall Boundary ◽  
Turbine Cascades

Theories and experiments on secondary losses in axial turbine cascades without end clearance are reviewed. A formula is given which correlates the effect of blade loading on secondary losses more successfully than hitherto. However, it is also shown that secondary losses increase with upstream wall boundary layer thickness. Only a tentative expression for that effect can be suggested. In order to predict secondary losses reliably more must be known about these wall boundary layers.

Heat Transfer Around a Cylindrical Protuberance Mounted in a Plane Turbulent Boundary Layer

2005 ◽  
Vol 128 (2) ◽  
pp. 153-161 ◽  
Author(s):  
Takayuki Tsutsui ◽  
Masafumi Kawahara
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Aspect Ratio ◽  
Turbulent Boundary Layer ◽  
Layer Thickness ◽  
Boundary Layer Thickness ◽  
Reynolds Numbers ◽  
Heat Transfer Characteristics ◽  
Low Aspect Ratio ◽  
Maximum Value

Heat transfer characteristics around a low aspect ratio cylindrical protuberance placed in a turbulent boundary layer were investigated. The diameters of the protuberance, D, were 40 and 80mm, and the height to diameter aspect ratio H∕D ranged from 0.125 to 1.0. The Reynolds numbers based on D ranged from 1.1×104 to 1.1×105 and the thickness of the turbulent boundary layer at the protuberance location, δ, ranged from 26 to 120mm for these experiments. In this paper we detail the effects of the boundary layer thickness and the protuberance aspect ratio on heat transfer. The results revealed that the overall heat transfer for the cylindrical protuberance reaches a maximum value when H∕δ=0.24.

Subsonic Flow Over Open Cavities: Part 1 — Analytical Models and Numerical Investigations

2016 ◽  
Author(s):  
Weidong Shao ◽  
Jun Li
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Shear Layer ◽  
Layer Thickness ◽  
Boundary Layer Thickness ◽  
Feedback Loop ◽  
Subsonic Flow ◽  
Free Stream ◽  
Open Cavities ◽  
Free Stream Mach Number

The aeroacoustical oscillation and acoustic field generated by subsonic flow grazing over open cavities has been investigated analytically and numerically. The tone generation mechanism is elucidated with an analytical model based on the coupling between shear layer instabilities and acoustic feedback loop. The near field turbulent flow is obtained using two-dimensional Large Eddy Simulation (LES). A special mesh is used to absorb propagating disturbances and prevent spurious numerical reflections. Comparisons with available experimental data demonstrate good agreement in both the frequency and amplitude of the aeroacoustical oscillation. The physical phenomenon of the noise generated by the feedback loop is discussed. The correlation analysis of primitive variables is also made to clarify the characteristics of wave propagation in space and time. The effects of free-stream Mach number and boundary layer thickness on pressure fluctuations within the cavity and the nature of the noise radiated to the far field are examined in detail. As free-stream Mach number increases velocity fluctuations and mass flux into the cavity increase, but the resonant Strouhal numbers slightly decrease. Both the resonant Strouhal numbers and sound pressure levels decrease with the increase of boundary layer thickness. Results indicate that the instability of the shear layer dominates both the frequency and amplitude of the aeroacoustical oscillation.

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