scholarly journals Periodicity and Repetivity of Unsteady Measurements of an Annular Turbine Cascade at Off Design Flow Conditions

1993 ◽  
Author(s):  
A. Bölcs ◽  
H. Körbächer
Keyword(s):  
Steady State ◽  
Test Data ◽  
Separation Bubble ◽  
Incidence Angle ◽  
Leading Edge ◽  
Design Flow ◽  
Turbine Cascade ◽  
Flow Conditions ◽  
Suction Surface ◽  
Shock Region

A two-dimensional section of a gas turbine cascade has been investigated experimentally in an annular non-rotating cascade facility as regards to its steady-state and time-dependent aerodynamic characteristics at off-design flow conditions. The blades vibrated in the first traveling wave bending mode. Steady-state and unsteady data were obtained for an off-design incidence angle of about 22° and for an isentropic outlet Mach number of M2s=1.19. At this flow condition, a separation bubble was present on the suction surface close to the leading edge. A shock appeared at trans- and supersonic outlet flow conditions on the suction surface. The data showed high unsteady loads close to the leading edge and in the shock region. It was found that the steady and the unsteady pressures in the shock region on the blade surface seemed to be very sensitive to small changes in the flow conditions. The periodicity and repetitivity of the steady and the unsteady pressures (σ=180°) was checked at several circumferential channel positions. This was done to figure out to which extend test data obtained in an annular ring channel can serve as a basis for the comparison with numerically obtained data. The aim of this paper is to show where problems may arise when comparing calculated results with test data.

Studies on Two-Dimensional Contouring of High-Lift Turbine Airfoil Suction Surface as Separation-Control Device: Separation Suppression Under Steady-State Flow Conditions

2011 ◽  
Author(s):  
Ken-ichi Funazaki ◽  
Nozomi Tanaka ◽  
Takahiro Shiba ◽  
Haruyuki Tanimitsu ◽  
Masaaki Hamabe
Keyword(s):  
Steady State ◽  
Separation Bubble ◽  
Control Device ◽  
Surface Boundary ◽  
Two Dimensional ◽  
Flow Conditions ◽  
Steady State Flow ◽  
Suction Surface ◽  
Airfoil Surface ◽  
State Flow

The study the present authors have been working on is to develop a new method to increase aerodynamic loading of low-pressure turbine airfoils for modern aeroengines to a great extent, which is to achieve drastic reduction of their airfoil counts. For this purpose, this study proposes two-dimensional contouring of the airfoil suction surface as a device to suppress the separation bubble that causes large aerodynamic loss, especially at low Reynolds number condition. The main objective of this paper is to show how and to what extent the surface contouring without any other disturbances affects the suction surface boundary layer accompanying separation bubble. For comparison, rather conventional tripping wire technique is also employed as “local 2D surface contouring” to generate flow disturbances in order to suppress the separation bubble. All measurements are carried out under steady-state flow conditions with low freestream turbulence. It turns out from the detailed experiments and LES analysis that the newly proposed two-dimensional contouring of the airfoil surface can effectively suppress the separation bubble, resulting in significant improvement of cascade aerodynamic performance.

Analysis of a Tilted Turbine Housing Volute Design Under Pulsating Inlet Conditions

2017 ◽  
Author(s):  
Samuel P. Lee ◽  
Martyn L. Jupp ◽  
Ambrose K. Nickson ◽  
John M. Allport
Keyword(s):  
Steady State ◽  
Incidence Angle ◽  
Leading Edge ◽  
Pulsating Flow ◽  
Flow Conditions ◽  
Low Velocity ◽  
Mixed Flow ◽  
Turbine Wheel ◽  
Housing Design ◽  
Inlet Conditions

Radial inflow turbines are widely used in the automotive turbocharger industry due to the greater amount of work that can be extracted per stage and their ease of manufacture compared with equivalent axial designs [1]. The current industry trend towards downsized engines for lower emissions has driven research to focus on improving turbine technologies for greater aero-thermal efficiency. Consequently, mixed flow turbines have recently received significant interest due to a number of potential performance benefits over their radial counterparts, including reduced inertia and improved performance at low velocity ratios. This paper investigates the performance of a tilted volute design compared with that of a radial design, under steady state and pulsating flow conditions. The tilted volute design was introduced in an attempt to improve inlet flow conditions of a mixed flow turbine wheel and hence improve performance. The investigation is entirely computational and the approach used was carefully validated against gas stand test results. The results of the study show that under steady state conditions the tilted volute design resulted in stage efficiency improvements of up to 1.64%. Under pulsating flow conditions, the tilted housing design resulted in a reduction in incidence angle and a maximum cycle averaged rotor efficiency improvement of 1.49% while the stage efficiencies resulted in a 1.23% increase. To assess the loss mechanisms within the rotor, the entropy flux generation through the blade passage was calculated. The tilted housing design resulted in reductions in leading edge suction and shroud surface separation resulting in the improved efficiency as observed.

The Effect of Reduced Frequency on Transition and Separation at the Leading Edge of a Compressor Stator

2012 ◽  
Author(s):  
Samuel C. T. Perkins ◽  
Alan D. Henderson
Keyword(s):  
Pressure Gradient ◽  
Steady Flow ◽  
Separation Bubble ◽  
Leading Edge ◽  
Flow Conditions ◽  
Suction Surface ◽  
Flow Angle ◽  
Pressure Surface ◽  
Reduced Frequency ◽  
Wake Passing

Studies on the effects of stator reduced frequency in low pressure turbines have shown that periodic wake-induced unsteadiness can increase steady flow circulation by as much as 15% and reduce losses compared to a steady flow datum. A large separation bubble downstream of peak suction that formed under steady flow conditions was periodically suppressed by wake passing events, resulting in significantly reduced losses within the boundary layer. This research extends this concept to a controlled diffusion compressor stator blade with a circular arc leading edge. The blade was placed inside a large scale, two-dimensional, cascade with a rotating bar mechanism used to simulate an upstream rotor blade row. The blade profile has been shown to experience leading edge separations and subsequent transition on both the pressure and suction surfaces due to a velocity overspeed caused by discontinuities in surface curvature. Testing was carried out at reduced frequencies of 0.47, 0.94 and 1.88 at the design inlet flow angle 45.5° and Reynolds number based on chord of 230,000. The freestream turbulence intensity was 4.0%. A range of experimental measurements were used to look at the blade’s performance: high resolution time-averaged blade surface static pressure measurements, inlet and exit 3-hole probe traverses and instantaneous, ensemble averaged and time average surface mounted hot-film measurements for the calculation of turbulent intermittency and quasi wall-shear stress. Results showed that increasing the stator reduced frequency from, 0–1.88, increased the overall blade pressure loss. The losses generated by the pressure surface and suction surface differed significantly and are affected very differently. The pressure surface demonstrated a clear reduction in loss with an increase in reduced frequency whereas the opposite trend was seen on the suction surface. Wake-induced turbulent strips suppressed the formation of leading edge separation bubbles that formed under steady flow conditions and in between wake passing events. Wake-induced turbulent strips reduced in width and level of turbulent intermittency through the favorable pressure gradients leading to peak suction and grew in the adverse pressure gradient of the velocity overspeed. The flow between wake-induced turbulent strips partially relaminarised through the favorable pressure gradient leading to peak suction.

scholarly journals Oscillating Cascade Aerodynamics at Large Mean Incidence

1998 ◽  
Vol 120 (1) ◽  
pp. 122-130 ◽  
Author(s):  
D. H. Buffum ◽  
V. R. Capece ◽  
A. J. King ◽  
Y. M. EL-Aini
Keyword(s):  
Separation Bubble ◽  
Incidence Angle ◽  
Separated Flow ◽  
Leading Edge ◽  
Pressure Measurements ◽  
Suction Surface ◽  
Airfoil Surface ◽  
Attached Flow ◽  
Blade Tip ◽  
Fan Blade

The aerodynamics of a cascade of airfoils oscillating in torsion about the midchord is investigated experimentally at a large mean incidence angle and, for reference, at a low mean incidence angle. The airfoil section is representative of a modern, low-aspect-ratio, fan blade tip section. Time-dependent airfoil surface pressure measurements were made for reduced frequencies of up to 1.2 for out-of-phase oscillations at a Mach number of 0.5 and chordal incidence angles of 0 and 10 deg; the Reynolds number was 0.9 × 106. For the 10 deg chordal incidence angle, a separation bubble formed at the leading edge of the suction surface. The separated flow field was found to have a dramatic effect on the chordwise distribution of the unsteady pressure. In this region, substantial deviations from the attached flow data were found, with the deviations becoming less apparent in the aft region of the airfoil for all reduced frequencies. In particular, near the leading edge the separated flow had a strong destabilizing influence while the attached flow had a strong stabilizing influence.

scholarly journals Oscillating Cascade Aerodynamics at Large Mean Incidence

1996 ◽  
Author(s):  
Daniel H. Buffum ◽  
Vincent R. Capece ◽  
Aaron J. King ◽  
Yehia M. El-Aini
Keyword(s):  
Separation Bubble ◽  
Incidence Angle ◽  
Separated Flow ◽  
Leading Edge ◽  
Pressure Measurements ◽  
Suction Surface ◽  
Airfoil Surface ◽  
Attached Flow ◽  
Blade Tip ◽  
Fan Blade

The aerodynamics of a cascade of airfoils oscillating in torsion about the midchord is investigated experimentally at a large mean incidence angle and, for reference, at a low mean incidence angle. The airfoil section is representative of a modern, low aspect ratio, fan blade tip section. Time-dependent airfoil surface pressure measurements were made for reduced frequencies of up to 1.2 for out-of-phase oscillations at a Mach number of 0.5 and chordal incidence angles of 0° and 10°; the Reynolds number was 0.9×106. For the 10° chordal incidence angle, a separation bubble formed at the leading edge of the suction surface. The separated flow field was found to have a dramatic effect on the chordwise distribution of the unsteady pressure. In this region, substantial deviations from the attached flow data were found with the deviations becoming less apparent in the aft region of the airfoil for all reduced frequencies. In particular, near the leading edge the separated flow had a strong destabilizing influence while the attached flow had a strong stabilizing influence.

The Measurement of Boundary Layers on a Compressor Blade in Cascade: Part 2—Suction Surface Boundary Layers

1988 ◽  
Vol 110 (1) ◽  
pp. 138-145 ◽  
Author(s):  
S. Deutsch ◽  
W. C. Zierke
Keyword(s):  
Boundary Layers ◽  
Separation Bubble ◽  
Incidence Angle ◽  
Outer Region ◽  
Leading Edge ◽  
Compressor Blade ◽  
Surface Boundary ◽  
Suction Surface ◽  
Circular Arc ◽  
Edge Separation

Using the facility described in Part 1 [29], eleven detailed velocity and turbulence intensity profiles are obtained on the suction surface of a double circular arc blade in cascade. At the measured incidence angle of 5 deg, transition through a leading edge separation bubble occurs before 2.6 percent chord. A continuing recovery from this leading edge separation is apparent in the measured boundary layer profiles at 2.6 and 7.6 percent chord. Recovery appears to be complete by 12.7 percent chord. The data then illustrate the evolution of the nonequilibrium turbulent boundary layers as they approach a second region of separation. Following the criteria established by Simpson et al. [1], we find that intermittent separation occurs near 60 percent chord while detachment occurs at 84.2 percent chord. Comparison between the measured profiles and the sublimation visualization studies indicates that the flow visualization is signaling the location of incipient detachment (1 percent instantaneous backflow). Measured profiles are also considered in light of similarity techniques for boundary layers approaching separation. Outer region similarity is shown to vanish for profiles downstream of detachment.

Experimental Investigation on the Annular Sector Cascade of a High Endwall-Angle Turbine

2016 ◽  
Author(s):  
Weiliang Fu ◽  
Jie Gao ◽  
Chen Liang ◽  
Fukai Wang ◽  
Qun Zheng ◽  
...  
Keyword(s):  
Mach Number ◽  
Static Pressure ◽  
Incidence Angle ◽  
Pressure Sensors ◽  
Leading Edge ◽  
Control Program ◽  
Turbine Cascade ◽  
Cascade Flow ◽  
Annular Sector ◽  
The Way

The flow in high endwall-angle turbine is complex, and it is different from the ordinary turbine flow in characteristics. In order to study the flow field characteristics of high endwall-angle turbines, the annular sector cascade experimental study of high endwall-angle turbines is carried out. The blade is studied experimentally in the form of annular sector cascade. The cascade includes 7 blades, and makes up 6 flow passages, in order to simulate full cascade flow. The experimental Mach number is adjusted by the way of changing inlet total pressure, and the Mach number influence (0.7, 0.8 and 0.9) on annular sector cascade flow is studied. Based on it, the inlet incidence angle (−15°, −7.5°, 0°, 7.5° and 15° )is changed with the way of changing sector straight pipes upstream of the cascade, and its influence on turbine flow fields is studied at the Mach number of 0.8. Here, five-hole probes are used to measure aerodynamic parameters distributions downstream of the cascade, and static pressure taps are positioned on the blade surface to measure surface static pressure distribution. The auto-traversing system and pressure sensors were operated by a self-compiled program based control program. The results indicate that there are two passage vortices inside the turbine cascade flow passage under the high Mach number condition, and the passage vortex near the high endwall-angle region is bigger. As Mach number increases, the passage vortices inside turbine cascade passage will become strong, and moves towards the blade mid-span. Besides, it is shown that the way of changing sector straight pipes can achieve the variation of inlet incidence angles. And, the blade profile with big leading-edge radius has good design and off-design performance. Detailed results and analyses are presented in the paper.

Large Eddy Simulation of Wake-Shear Layer Interactions Over a Multi-Element Aerofoil

2020 ◽  
Author(s):  
Souvik Naskar ◽  
S. Sarkar
Keyword(s):  
Large Eddy Simulation ◽  
Shear Layer ◽  
Separation Bubble ◽  
Turbine Blades ◽  
Leading Edge ◽  
Aerodynamic Characteristics ◽  
Suction Surface ◽  
Free Stream Turbulence ◽  
Eddy Simulation ◽  
Large Eddy

Abstract Modern commercial airliners use multi-element aerofoils to enhance take-off and landing performance. Further, multielement aerofoil configurations have been shown to improve the aerodynamic characteristics of wind turbines. In the present study, high resolution Large Eddy Simulation (LES) is used to explore the low Reynolds Number (Re = 0.832 × 104) aerodynamics of a 30P30N multi-element aerofoil at an angle of attack, α = 4°. In the present simulation, wake shed from a leading edge element or slat is found to interact with the separated shear layer developing over the suction surface of the main wing. High receptivity of shear layer via amplification of free-stream turbulence leads to rollup and breakdown, forming a large separation bubble. A transient growth of fluctuations is observed in the first half of the separation bubble, where levels of turbulence becomes maximum near the reattachment and then decay depicting saturation of turbulence. Results of the present LES are found to be in close agreement with the experiment depicting high vortical activity in the outer layer. Some features of the flow field here are similar to those occur due to interactions of passing wake and the separated boundary layer on the suction surface of high lift low pressure turbine blades.

Numerical Investigation of the Effect of Inlet Flow Angle on Film Cooling Performance for the Blade Tip With Suction Surface Squealer Rim

2019 ◽  
Author(s):  
Zhiqiang Yu ◽  
Jianjun Liu ◽  
Chen Li ◽  
Baitao An
Keyword(s):  
Mass Flow ◽  
Film Cooling ◽  
Incidence Angle ◽  
Leading Edge ◽  
Flow Ratio ◽  
Suction Surface ◽  
Cooling Performance ◽  
Tip Leakage ◽  
Pressure Surface ◽  
Mass Flow Ratio

Abstract Numerical investigations have been performed to study the effect of incidence angle on the aerodynamic and film cooling performance for the suction surface squealer tip with different film-hole arrangements at τ = 1.5% and BR = 1.0. Meanwhile, the full squealer tip as baseline is also investigated. Three incidence angles at design condition (0 deg) and off-design conditions (± 7 deg) are investigated. The suction surface, pressure surface, and the camber line have seven holes each, with an extra hole right at the leading edge. The Mach number at the cascade inlet and outlet are 0.24 and 0.52, respectively. The results show that the incidence angle has a significant effect on the tip leakage flow characteristics and coolant flow direction. The film cooling effectiveness distribution is altered, especially for the film holes near the leading edge. When the incidence angle changes from +7 deg to 0 and −7 deg, the ‘re-attachment line’ moves downstream and the total tip leakage mass flow ratio decreases, but the suction surface tip leakage mass flow ratio near leading edge increases. In general, the total tip leakage mass flow ratio for suction surface squealer tip is 1% greater than that for full squealer tip at the same incidence angle. The total pressure loss coefficient of suction surface squealer tip is larger than that for full squealer tip. The full squealer tip with film holes near suction surface and the suction surface squealer tip with film hole along camber line show high film cooling performance, and the area averaged film cooling effectiveness at positive incidence angle +7 deg is higher than that at 0 and −7 deg. The coolant discharged from film holes near pressure surface only cools narrow region near pressure surface.

Study on Leakage Loss Control Method of Circumferential Bending Clearance

2019 ◽  
Author(s):  
Shuai Jiang ◽  
Fu Chen ◽  
Jianyang Yu ◽  
Shaowen Chen ◽  
Yanping Song
Keyword(s):  
Control Method ◽  
Incidence Angle ◽  
Leading Edge ◽  
Blocking Effect ◽  
Field Analysis ◽  
Leakage Flow ◽  
Turbine Cascade ◽  
Tip Leakage Flow ◽  
Control Effect ◽  
Tip Leakage

Abstract The concept of circumferential bending clearance based on Gauss Bimodal Function is proposed to suppress tip leakage flow (TLF) in a highly-loaded turbine cascade. In this method, a new vortex (BV) can be induced to mix with TLV in the middle of tip region and block the development of tip leakage vortex (TLV). Since the blocking effect divides the TLV into two parts, the tip leakage rate and loss of TLF can be reduced significantly. In order to reveal the mechanisms of blocking effect on leakage flow and its influencing factors, the research numerically investigates the effects of environmental conditions on the TLF development in a turbine cascade. The flow field analysis of the optimal bending clearance is in the first place, and then the effects of clearance heights (δ) and incidence angles (α) on the TLF characteristic and loss are investigated respectively. Results indicate that the blocking effect has a close relationship with the TLF characteristic, which can be divided into the BV migration, TLV-2 location and blocking loss. The nearer distance to the leading edge (LE) and farther distance to the suction side (SS) of BV means a less loss of TLF in bending clearance cases. The further distance away from blade tip and SS of TLV-2 means a larger-scale vortex with more loss. The additional loss in blocking region expands constantly with the increase of clearance height and incidence angle. The bending clearance has limited control effect on TLF with the variation of clearance height, especially the loss increases in Case 2%H. However, it has a strong adaptability with the change of incidence angle, the relative total pressure loss drops up to 16% in Case −5°.

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