The Present Challenge of Transonic Compressor Blade Design

2019 ◽  
Vol 141 (9) ◽  
Author(s):  
Alexander Hergt ◽  
J. Klinner ◽  
J. Wellner ◽  
C. Willert ◽  
S. Grund ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Shock Waves ◽  
Unsteady Flow ◽  
Flow Behavior ◽  
Complex Structure ◽  
Compressor Cascade ◽  
Transonic Compressor ◽  
Flow Phenomena ◽  
Transonic Cascade

The flow through a transonic compressor cascade shows a very complex structure due to the occurring shock waves. In addition, the interaction of these shock waves with the blade boundary layer inherently leads to a very unsteady flow behavior. The aim of the current investigation is to quantify this behavior and its influence on the cascade performance as well as to describe the occurring transonic flow phenomena in detail. Therefore, an extensive experimental investigation of the flow in a transonic compressor cascade has been conducted within the transonic cascade wind tunnel of DLR Institute of Propulsion Technology at Cologne. In this process, the flow phenomena were thoroughly examined for an inflow Mach number of 1.21. The experiments investigate both the laminar and the turbulent shock wave boundary layer interaction within the blade passage and the resulting unsteady behavior. The experiments show a fluctuation range of the passage shock wave of about 10% chord for both cases, which is directly linked with a change of the inflow angle and of the operating point of the cascade. Thereafter, Reynolds-averaged Navier–Stokes (RANS) simulations have been performed aiming at the verification of the reproducibility of the experimentally examined flow behavior. Here, it is observed that the dominant flow effects are not reproduced by a steady numerical simulation. Therefore, a further unsteady simulation has been carried out to capture the unsteady flow behavior. The results from this simulation show that the fluctuation of the passage shock wave can be reproduced but not in the correct magnitude. This leads to a remaining weak point within the design process of transonic compressor blades because the working range will be overpredicted. The resulting conclusion of this study is that the use of scale-resolving methods such as LES or the application of DNS is necessary to correctly predict unsteadiness of the transonic cascade flow and its impact on the cascade performance.

The Present Challenge of Transonic Compressor Blade Design

2018 ◽  
Author(s):  
A. Hergt ◽  
J. Klinner ◽  
J. Wellner ◽  
C. Willert ◽  
S. Grund ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Shock Waves ◽  
Unsteady Flow ◽  
Flow Behavior ◽  
Flow Behaviour ◽  
Compressor Cascade ◽  
Transonic Compressor ◽  
Flow Phenomena ◽  
Transonic Cascade

The flow through a transonic compressor cascade shows a very complex structure due to the occuring shock waves. In addition, the interaction of these shock waves with the blade boundary layer inherently leads to a very unsteady flow behaviour. The aim of the current investigation is to quantify this behaviour and its influence on the cascade performance as well as to describe the occuring transonic flow phenomena in detail. Therefore, an extensive experimental investigation of the flow in a transonic compressor cascade has been conducted within the transonic cascade wind tunnel of DLR at Cologne. In this process, the flow phenomena were thoroughly examined for an inflow Mach number of 1.21. The experiments investigate both, the laminar as well as the turbulent shock wave boundary layer interaction within the blade passage and the resulting unsteady behaviour. The experiments show a fluctuation range of the passage shock wave of about 10 percent chord for both cases, which is directly linked with a change of the inflow angle and of the operating point of the cascade. Thereafter, RANS simulations have been performed aiming at the verification of the reproducibility of the experimentally examined flow behavior. Here it is observed that the dominant flow effects are not reproduced by a steady numerical simulation. Therefore, a further unsteady simulation has been carried out in order to capture the unsteady flow behaviour. The results from this simulation show that the fluctuation of the passage shock wave can be reproduced but not in the correct magnitude. This leads to a remaining weak point within the design process of transonic compressor blades, because the working range will be overpredicted. The resulting conclusion of the study is that the use of scale resolving methods such as LES or the application of DNS is necessary to correctly predict unsteadiness of the transonic cascade flow and its impact on the cascade performance.

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.

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.

scholarly journals Aerodynamic Inverse Design of Transonic Compressor Cascades with Stabilizing Elastic Surface Algorithm

2021 ◽  
Vol 11 (11) ◽  
pp. 4845
Author(s):  
Mohammad Hossein Noorsalehi ◽  
Mahdi Nili-Ahmadabadi ◽  
Seyed Hossein Nasrazadani ◽  
Kyung Chun Kim
Keyword(s):  
Design Method ◽  
Inverse Design ◽  
Normal Shock ◽  
Compressor Cascade ◽  
Elastic Surface ◽  
Transonic Compressor ◽  
Pressure Distributions ◽  
The Difference ◽  
Inverse Design Method ◽  
Transonic Cascade

The upgraded elastic surface algorithm (UESA) is a physical inverse design method that was recently developed for a compressor cascade with double-circular-arc blades. In this method, the blade walls are modeled as elastic Timoshenko beams that smoothly deform because of the difference between the target and current pressure distributions. Nevertheless, the UESA is completely unstable for a compressor cascade with an intense normal shock, which causes a divergence due to the high pressure difference near the shock and the displacement of shock during the geometry corrections. In this study, the UESA was stabilized for the inverse design of a compressor cascade with normal shock, with no geometrical filtration. In the new version of this method, a distribution for the elastic modulus along the Timoshenko beam was chosen to increase its stiffness near the normal shock and to control the high deformations and oscillations in this region. Furthermore, to prevent surface oscillations, nodes need to be constrained to move perpendicularly to the chord line. With these modifications, the instability and oscillation were removed through the shape modification process. Two design cases were examined to evaluate the method for a transonic cascade with normal shock. The method was also capable of finding a physical pressure distribution that was nearest to the target one.

scholarly journals Effects of Tip Leakage Flow on the Aerodynamic Performance and Stability of an Axial-Flow Transonic Compressor Stage

Energies ◽  
2021 ◽  
Vol 14 (14) ◽  
pp. 4168
Author(s):  
Botao Zhang ◽  
Xiaochen Mao ◽  
Xiaoxiong Wu ◽  
Bo Liu
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Axial Flow ◽  
Leading Edge ◽  
Rotating Stall ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Tip Leakage ◽  
Transonic Compressor

To explain the effect of tip leakage flow on the performance of an axial-flow transonic compressor, the compressors with different rotor tip clearances were studied numerically. The results show that as the rotor tip clearance increases, the leakage flow intensity is increased, the shock wave position is moved backward, and the interaction between the tip leakage vortex and shock wave is intensified, while that between the boundary layer and shock wave is weakened. Most of all, the stall mechanisms of the compressors with varying rotor tip clearances are different. The clearance leakage flow is the main cause of the rotating stall under large rotor tip clearance. However, the stall form for the compressor with half of the designed tip clearance is caused by the joint action of the rotor tip stall caused by the leakage flow spillage at the blade leading edge and the whole blade span stall caused by the separation of the boundary layer of the rotor and the stator passage. Within the investigated varied range, when the rotor tip clearance size is half of the design, the compressor performance is improved best, and the peak efficiency and stall margin are increased by 0.2% and 3.5%, respectively.

Characterization of Shock Wave–Endwall Boundary Layer Interactions in a Transonic Compressor Rotor

1988 ◽  
Vol 110 (3) ◽  
pp. 386-392 ◽  
Author(s):  
D. C. Rabe ◽  
A. J. Wennerstrom ◽  
W. F. O’Brien
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Air Force ◽  
Leading Edge ◽  
Transonic Compressor ◽  
Laser Measurements ◽  
Compressor Rotor ◽  
Boundary Layer Interaction ◽  
Air Force Base

The passage shock wave–endwall boundary layer interaction in a transonic compressor was investigated with a laser transit anemometer. The transonic compressor used in this investigation was developed by the General Electric Company under contract to the Air Force. The compressor testing was conducted in the Compressor Research Facility at Wright-Patterson Air Force Base, OH. Laser measurements were made in two blade passages at seven axial locations from 10 percent of the axial blade chord in front of the leading edge to 30 percent of the axial blade chord into the blade passage. At three of these axial locations, laser traverses were taken at different radial immersions. A total of 27 different locations were traversed circumferentially. The measurements reveal that the endwall boundary layer in this region is separated from the core flow by what appears to be a shear layer where the passage shock wave and all ordered flow seem to end abruptly.

Effect of Shock Wave Behavior on Unsteady Aerodynamic Characteristics of Oscillating Transonic Compressor Cascade

2021 ◽  
Author(s):  
Jiuliang Gan ◽  
Toshinori Watanabe ◽  
Takehiro Himeno
Keyword(s):  
Shock Wave ◽  
Aerodynamic Force ◽  
Aerodynamic Characteristics ◽  
Fast Response ◽  
Influence Coefficient ◽  
Compressor Cascade ◽  
Blade Vibration ◽  
Unsteady Pressure ◽  
Transonic Compressor ◽  
Unsteady Aerodynamic

Abstract The unsteady behavior of the shock wave was studied in an oscillating transonic compressor cascade. The experimental measurement and corresponding numerical simulation were conducted on the cascade with different shock patterns based on influence coefficient method. The unsteady pressure distribution on blade surface was measured with fast-response pressure-sensitive paint (PSP) to capture the unsteady aerodynamic force as well as the shock wave movement. It was found that the movement of shock waves in the neighboring flow passages of the oscillating blade was almost anti-phase between the two shock patterns, namely, the double shocks pattern and the merged shock pattern. It was also found that the amplitude of the unsteady pressure caused by the passage shock wave was very large under the merged shock pattern compared with the double shocks pattern. The stability of blade vibration was also analyzed for both shock patterns including 3-D flow effect. These findings were thought to shed light on the fundamental understanding of the unsteady aerodynamic characteristics of oscillating cascade caused by the shock wave behavior.

Transition effect on aerodynamic performance of compressor profile

2020 ◽  
Vol 92 (4) ◽  
pp. 611-620
Author(s):  
Ryszard Szwaba ◽  
Piotr Kaczyński ◽  
Piotr Doerffer
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Aerodynamic Performance ◽  
Negative Influence ◽  
Good Choice ◽  
Boundary Layer Transition ◽  
Roughness Height ◽  
Compressor Cascade ◽  
Layer Transition ◽  
Content Type

Purpose The purpose of this paper is to study experimentally the effect of transition and also the roughness height on the flow structure of the shock wave boundary layer interaction in the blades passage of a compressor cascade. Design/methodology/approach A model of a turbine compressor passage was designed and assembled in a transonic wind tunnel. In the experiment, the distributed roughness with different heights and locations was used to induce transition upstream of the shock wave. Findings Recommendation regarding the roughness parameters for the application depends on what is more important as goal, whether the reduction of losses or unsteadiness. In case if more important are the losses reduction, a good choice for the roughness location seems to be the one close to the shock wave position. Research limitations/implications The knowledge gained by this paper will enable the implementation of an effective laminar flow technology for engines in which the interaction of a laminar boundary layer with a shock wave takes place in the propulsion system and causes severe problems. Originality/value The paper focuses on the influence of the boundary layer transition induced by different roughness values and locations on aerodynamic performance of a compressor cascade. Very valuable results were obtained in the roughness application for the boundary layer transition control, demonstrating a positive effect in changing the nature of the interaction and also some negative influence in case of oversized roughness height, which cannot be found in the existing literature.

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