Dynamics of Shock Waves Interacting With Laminar Separated Transonic Turbine Flow Investigated by High-Speed Schlieren and Surface Hot-Film Sensors

2020 ◽  
Author(s):  
Marcel Börner ◽  
Reinhard Niehuis
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
High Speed ◽  
Separation Bubble ◽  
Suction Side ◽  
Normal Shock ◽  
Experimental Investigations ◽  
Low Frequencies ◽  
Hot Film ◽  
Schlieren Images

Abstract The results presented in this paper are based on experimental investigations on a generic transonic low pressure turbine profile at high subsonic exit Mach numbers. Here, the flow on the suction side reaches a maximum isentropic Mach number of approximately 1.2 and features a large separation bubble in a transonic flow regime characterized by Surface Hot-Film measurements. The measurements are supplemented by Schlieren images recorded with a high-speed camera at 19.2 kHz. A highly unsteady normal shock wave on the suction side is observable upstream of the trailing edge. It is interacting with laminar separated flow which is rarely documented in literature. The interaction of the normal shock with the boundary layer flow seems to amplifies the ongoing transition process over the separation bubble and the flow reattaches shortly downstream. A statistical analysis of the Schlieren images reveals characteristic low frequencies of the shock wave motions and a pulsation of the separation bubble. Additionally, the statistical information of the time-dependent signal from the Surface Hot-Film sensors demonstrate the instabilities influencing the boundary layer linked to the unsteadiness in the main flow.

DYNAMICS OF SHOCK WAVES INTERACTING WITH LAMINAR SEPARATED TRANSONIC TURBINE FLOW INVESTIGATED BY HIGH-SPEED SCHLIEREN AND SURFACE HOT-FILM SENSORS

2021 ◽  
pp. 1-12
Author(s):  
Marcel Börner ◽  
Reinhard Niehuis
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
High Speed ◽  
Separation Bubble ◽  
Suction Side ◽  
Normal Shock ◽  
Experimental Investigations ◽  
Low Frequencies ◽  
Hot Film ◽  
Schlieren Images

Abstract The results presented in this paper are based on experimental investigations on a generic transonic low pressure turbine profile at high subsonic exit Mach numbers. Here, the flow on the suction side reaches a maximum isentropic Mach number of approximately 1.2 and features a large separation bubble in a transonic flow regime characterized by Surface Hot-Film measurements. The measurements are supplemented by Schlieren images recorded with a high-speed camera at 19:2 kHz. A highly unsteady normal shock wave on the suction side is observable upstream of the trailing edge. It is interacting with laminar separated flow which is rarely documented in literature. The interaction of the normal shock with the boundary layer flow seems to amplifies the ongoing transition process over the separation bubble and the flow reattaches shortly downstream. A statistical analysis of the Schlieren images reveals characteristic low frequencies of the shock wave motions and a pulsation of the separation bubble. Additionally, the statistical information of the time-dependent signal from the Surface Hot-Film sensors demonstrate the instabilities influencing the boundary layer linked to the unsteadiness in the main flow.

ON THE IMPORTANCE OF TRANSITION CONTROL AT TRANSONIC COMPRESSOR BLADES

2021 ◽  
pp. 1-48
Author(s):  
Alexander Hergt ◽  
Joachim Klinner ◽  
Sebastian Grund ◽  
Chris Willert ◽  
Wolfgang Steinert ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
High Speed ◽  
Separation Bubble ◽  
Experimental Investigations ◽  
Compressor Cascade ◽  
Transition Control ◽  
Transonic Compressor ◽  
Air Jet ◽  
Loss Level

Abstract The flow through a transonic compressor cascade is characterized by high unsteadiness and a high loss level. In the case of a laminar shock wave boundary layer interaction the loss level is higher due to the occurrence of a laminar separation bubble below the shock wave compared to the shock wave interaction with a turbulent boundary layer. In addition, the oscillation of the shock position in both cases influences the working range concerning the point of stall onset as well as leading to an unsteady interaction with the blade, called buffeting. The reduction of losses and of unsteadiness in the shock wave oscillation, connected to a decrease of the blade buffeting effect, are the aims of the current investigation. Therefore, experimental investigations using a roughness patch as well as air jet vortex generators in order to control the transition in a transonic compressor cascade have been conducted at the transonic cascade wind tunnel of DLR at Cologne. At an inflow Mach number of 1.21 a loss reduction for both transition control cases is achieved. In spite of a nearly uninfluenced fluctuation range of the passage shock wave compared to the reference cascade, the oscillation spectra of the transition control cases show a reduction of the shock movement amplitude at a frequency below 500 Hz and above 1 kHz. In the closing section of the paper a detailed discussion on the reasons for the resulting flow behaviour based on PIV and High Speed Shadowgraphy data is given.

Turbine Cascade Flow Control Using a “Wake Filling” Pulsed Plasma Actuator

2005 ◽  
Author(s):  
H. Perez-Blanco ◽  
Robert Van Dyken ◽  
Aaron Byerley ◽  
Tom McLaughlin
Keyword(s):  
Boundary Layer ◽  
Pressure Gradient ◽  
Laminar Boundary Layer ◽  
Separation Bubble ◽  
Adverse Pressure Gradient ◽  
Suction Side ◽  
Plasma Actuator ◽  
Pulsed Plasma ◽  
Power Cost ◽  
Hot Film

Separation bubbles in high-camber blades under part-load conditions have been addressed via continuous and pulsed jets, and also via plasma actuators. Numerous passive techniques have been employed as well. In this type of blades, the laminar boundary layer cannot overcome the adverse pressure gradient arising along the suction side, resulting on a separation bubble. When separation is abated, a common explanation is that kinetic energy added to the laminar boundary layer speeds up its transition to turbulent. In the present study, a plasma actuator installed in the trailing edge (i.e. “wake filling configuration”) of a cascade blade is used to excite the flow in pulsed and continuous ways. The pulsed excitation can be directed to the frequencies of the large coherent structures (LCS) of the flow, as obtained via a hot-film anemometer, or to much higher frequencies present in the suction-side boundary layer, as given in the literature. It is found that pulsed frequencies much higher than that of LCS reduce losses and improve turning angles further than frequencies close to those of LCS. With the plasma actuator 50% on time, good loss abatement is obtained. Larger “on time” values yield improvements, but with decreasing returns. Continuous high-frequency activation results in the largest loss reduction, at increased power cost. The effectiveness of high frequencies may be due to separation abatement via boundary layer excitation into transition, or may simply be due to the creation of a favorable pressure gradient that averts separation as the actuator ejects fluid downstream. Both possibilities are discussed in light of the experimental evidence.

Experimental Investigation of the Clocking Effect in a 1.5-Stage Axial Turbine—Part II: Unsteady Results and Boundary Layer Behavior

2009 ◽  
Vol 131 (2) ◽  
Author(s):  
Sven König ◽  
Bernd Stoffel ◽  
M. Taher Schobeiri
Keyword(s):  
Boundary Layer ◽  
Separation Bubble ◽  
Measurement Techniques ◽  
Leading Edge ◽  
Suction Side ◽  
Experimental Investigations ◽  
Stress Signal ◽  
Upstream Direction ◽  
Loss Parameter ◽  
Lp Turbine

Comprehensive experimental investigations were conducted to get deeper insight into the physics of stator clocking in turbomachines. Different measurement techniques were used to investigate the influence of varying clocking positions on the highly unsteady flow field in a 1.5-stage axial low-pressure (LP) turbine. A Reynolds number typical for LP turbines as well as a two-dimensional blade design were chosen. Stator 2 was developed as a high-lift profile with a separation bubble on the suction side. This paper presents the results that were obtained by means of unsteady x-wire measurements upstream and downstream of Stator 2 and surface mounted hot-film measurements on the Stator 2 suction side. It was found that for the case when the Stator 1 wakes impinge close to the leading edge of Stator 2 the interaction between the Stator 1 and the rotor vortical structures takes place in proximity of the Stator 2 boundary layer, which leads to a shift of the transition point in the upstream direction. The major loss parameter concerning the Stator 2 aerodynamic performance could be attributed to the strength of the periodic fluctuations within the Stator 2 suction side boundary layer. A phase shift in the quasiwall shear stress signal in the front region of the Stator 2 vane was observed for different clocking positions.

Aerodynamic Measurements on a Low Pressure Turbine Cascade With Different Levels of Distributed Roughness

2011 ◽  
Author(s):  
Marco Montis ◽  
Reinhard Niehuis ◽  
Andreas Fiala
Keyword(s):  
Boundary Layer ◽  
High Speed ◽  
Separation Bubble ◽  
Reference Conditions ◽  
Suction Side ◽  
Boundary Layer Separation ◽  
Low Pressure ◽  
Average Roughness ◽  
Low Pressure Turbine ◽  
High Reynolds

The effect of surface roughness on the aerodynamics of a highly loaded low-pressure turbine airfoil was investigated in a series of cascade tests conducted in a high speed facility. Profile loss and aerodynamic loading of three different surface roughnesses with a ratio of the centerline average roughness to the profile chord of 1.1 · 10−5, 7.1 · 10−5 and 29 · 10−5 were analysed. Tests were carried out under design outlet Mach number, outlet Reynolds number ranging from 5 · 104 to 7 · 105 and inlet turbulence level of 2.5% and 5%. The mid span flow field downstream of the cascade and the loading distribution on the profile were measured for each investigated operating point using a five hole probe and surface static pressure taps. Additional measurements with a hot wire probe in the profile boundary layer under reference conditions (Re2th = 2 · 105) were also conducted. Experimental results show a loss reduction for the highest roughness under reference conditions, due to the partial suppression of the separation bubble on the suction side of the profile. At high Reynolds numbers a massive boundary layer separation on the suction side is observed for the highest roughness, along with a large increase in total pressure loss. The middle roughness tested has no effect on the loading distribution as well as on the loss behaviour of the airfoil under all investigated flow conditions.

On the Applicability of a Spoked-Wheel Wake Generator for Clocking Investigations

2007 ◽  
Vol 129 (11) ◽  
pp. 1468-1477 ◽  
Author(s):  
Sven König ◽  
Bernd Stoffel
Keyword(s):  
Boundary Layer ◽  
Static Pressure ◽  
Separation Bubble ◽  
Measurement Techniques ◽  
Suction Side ◽  
Jet Engines ◽  
High Lift ◽  
Physical Mechanisms ◽  
Aero Engines ◽  
Hot Film

A comprehensive investigation was carried out using two different experimental setups: A 1.5-stage axial turbine and a simplified model, a “spoked-wheel” setup with a rotating wake generator consisting of cylindrical bars. The second stator of the turbine was designed at MTU Aero Engines as a high-lift profile with a Reynolds number typical for low-pressure turbines in jet engines. At design conditions, the flow on the stator 2 suction side features a pronounced separation bubble. To study the behavior of the stator 2 boundary layer and the interaction mechanisms between stator and rotor wakes, different measurement techniques were used: X-wire probes, five-hole probes, static pressure tappings, and surface mounted hot-film gauges. It was found that a rotating wake generator of the spoked-wheel type is not capable of resolving the relevant clocking mechanisms that occur in a real engine. However, such a simplified setup is useful to separate some of the physical mechanisms, and in case that the interaction of the stator 1 wakes with the stator 2 boundary layer is negligible, a spoked-wheel setup is well suited to simulate the influence of periodically incoming wakes on the transition behavior of stator 2.

Investigation of a Separated Boundary Layer and its Influence on Secondary Flow of a Transonic Turbine Profile

2014 ◽  
Author(s):  
Roland E. Brachmanski ◽  
Reinhard Niehuis ◽  
Arianna Bosco
Keyword(s):  
Boundary Layer ◽  
Secondary Flow ◽  
Turbine Blade ◽  
Flow Separation ◽  
High Speed ◽  
Separation Bubble ◽  
Armed Forces ◽  
Suction Side ◽  
Reynolds Numbers ◽  
Operation Conditions

Profile losses of the turbine blade and secondary flow losses are the main source of aerodynamic loss in a low pressure turbine. However, not much attention has been paid in the interaction between these two loss sources. This paper investigates the interaction mechanisms between a separated boundary layer on the suction side and the secondary flow in blade passages. The high speed cascade wind tunnel of the University of the Federal Armed Forces Germany has been used to achieve the required operation conditions, generating a flow separation on the suction side. The profile of this cascade has been chosen due to the flow separation behavior on the suction side of the blade at low Reynolds numbers. Different measurements techniques are conducted to further investigate the effects seen in CFD. The aim of this paper is to investigate the interaction phenomena between the secondary flow and a separation bubble at different Reynolds numbers. The development and change of the boundary layer in the axial and radial directions on the suction side of the turbine blade are presented and discussed. The results show discrepancies between the numerical prediction and the experimental data on the suction side of the blade rise as the effects of the secondary flow increase. Furthermore, the increasing influence of the radial pressure gradient of the secondary flow leads to a noticeable reduction in the length of the separation bubble close to the endwall region.

On the Importance of Transition Control at Transonic Compressor Blades

2019 ◽  
Author(s):  
A. Hergt ◽  
J. Klinner ◽  
S. Grund ◽  
C. Willert ◽  
W. Steinert ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
High Speed ◽  
Experimental Investigations ◽  
Compressor Cascade ◽  
Compressor Blades ◽  
Transition Control ◽  
Transonic Compressor ◽  
Air Jet ◽  
Loss Level

Abstract The flow through a transonic compressor cascade is characterized by high unsteadiness and a high loss level. This results from the shock waves in the blade cascade and their interaction with the blade suction side boundary layer. In the case of a laminar shock wave boundary layer interaction the loss level is higher due to the occurrence of a laminar separation bubble below the shock wave compared to the shock wave interaction with a turbulent boundary layer. In addition, the oscillation of the shock position in both cases influences the working range concerning the point of stall onset as well as leading to an unsteady interaction with the blade, called buffeting. The reduction of losses and of unsteadiness in the shock wave oscillation, connected to a decrease of the blade buffeting effect, are the aims of the current investigation. Therefore, experimental investigations using a roughness patch as well as air jet vortex generators in order to control the transition in a transonic compressor cascade have been conducted at the transonic cascade wind tunnel of DLR at Cologne. At an inflow Mach number of 1.21 a loss reduction for both transition control cases is achieved. In spite of a nearly uninfluenced fluctuation range of the passage shock wave compared to the reference cascade, the oscillation spectra of the transition control cases show a reduction of the shock movement amplitude at a frequency below 500 Hz and above 1 kHz. In the closing section of the paper a detailed discussion on the reasons for the resulting flow behaviour based on PIV and High Speed Shadowgraphy data is given. The resulting conclusion of the study is that the consideration of transition control at transonic compressor blades is very important in order to reduce losses and flow unsteadiness which directly influences blade buffeting and the numerical prediction quality of the stall onset.

Study of the Shock Wave–Turbulent Boundary Layer Interaction Using a 3D von Kármán Length Scale

2017 ◽  
Vol 18 (1) ◽  
pp. 57-66 ◽  
Author(s):  
Jing-Lei Xu ◽  
You-Fu Song ◽  
Yang Zhang ◽  
Jun-Qiang Bai
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Turbulence Model ◽  
High Speed ◽  
Compressible Flows ◽  
Separation Bubble ◽  
Turbulence Models ◽  
Length Scale ◽  
Von Karman ◽  
Von Kármán

AbstractTraditional turbulence models are initially formulated and calibrated under incompressible conditions. Thus, these models are always of low fidelity when extended to high speed, complex and compressible flows. In this work, a compressible von Kármán length scale is proposed for compressible flows considering the variable densities. The length scale is the ratio between the new vorticity and its gradient. The new length scale is actually based on phenomenological theory, which is then integrated into the KDO (turbulence Kinetic energy Dependent Only) turbulence model, arriving at a compressible model called CKDO (Compressible KDO). In the CKDO turbulence model, all the extra terms produced by compressibility are modeled as dissipation. Compression corners of 8, 16, 20 and 24 angles are studied within SST, SA, KDO and CKDO. These test cases are known as the typical shock wave–boundary layer interactions. The results show that the new length scale in CKDO is able to well capture the surface pressure and skin friction distributions. Besides, compared with the standard von Kármán length scale, the new length scale in CKDO can better capture the size and position of the separation bubble. With the increase of the corner angle, CKDO shows more prominent potential for describing compressible flows.

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