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

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.

Effects of Flow Distortion on Clarity of Interferometer Data

The goal of the paper is to propose modifications to tested flat cascades that will suppress partial distortions of acquired interferograms in the region of blade leading edges. This improvement will allow more accurate determination of actual flow incidence angles of tested blade cascades, in particular for transonic or supersonic inlet flows. Application of physical probes for such tasks is always in question in transonic or supersonic flows. The paper is composed of three main sections: (a) introduction of the test facility, (b) presentation of the problem with examples, and (c) description of the experimental work. Recommendations for future flat cascade investigations is presented in the paper. The first section is devoted to the introduction and description of the High-Speed Laboratory of the Institute of Thermomechanis of the Czech Academy of Sciences. Attention is paid to the unique large-scale interferometer which is one of the principal research instruments here and which is routinely used for investigations of transonic compressor and turbine cascades. The instrument capability is illustrated by a series of images showing evolution of a sonic line in a transonic cascade as a function of the increasing inlet Mach number. The reasoning for the proposed work is presented in the middle section. The major impetus for the work was to understand the observed discrepancies between schlieren and interferometer images while testing highly-loaded transonic compressor cascades. In particular, the main concern is the relatively wide region of increasing pressure in the shock vicinity recorded on interferograms versus sharp shock wave image visible on schlieren images. It was suggested that these discrepancies are caused by deformation of the shock-wave surface by the growth of secondary flow due to the tunnel endwall effects. It should be stressed here that the intention was not to investigate the pattern or the nature of the secondary flows rather. An idea behind this approach is to move the secondary flows out of the region of interferometer imaging. Finally, in the last section the results of the experiments carried out during the course of this work are presented. The experiments were designed to improve understanding of the origins of interferogram distortions. Further intention was to eliminate or at least lessen the level of interferogram distortions due to the combined effects of the boundary layer interaction and the corner-vortex flow. Wedges of a constant vertex angle of 15 deg of various plane shapes were inserted subsequently in supersonic flow of (Mach number 2) and interferograms of the resulting flow pattern were acquired. It was observed that decreasing the wedge span led to clearing the interferograms of the superimposed distortions. This confirmed the decisive role of the end wall effects on the quality of acquired results. The undistorted interferograms of the inlet flow in the region of the shock structure are needed to determine the actual angle of attack of the incoming flow onto the tested transonic cascade. Based on the presented results it is suggested for the future testing of flat cascades to modify the front part of the blades by appropriate side cut-offs to eliminate interferogram distortions.

The Present Challenge of Transonic Compressor Blade Design

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.

Supersonic Compressor Cascade Shape Optimization under Multiple Inlet Mach Operating Conditions

Transonic compressors are widely used today in propulsion and industrial applications thanks to their higher specific work compared to subsonic. In this work, the aerodynamic optimization of a two-dimensional Computational Fluid Dynamics (CFD) model of the transonic cascade ARL-SL19 is described. The validated computational model is used for a multi-objective optimization of the cascade at three different inlet Mach numbers using a genetic algorithm and an artificial neural network, with the aim of reducing total pressure loss and increasing maximum pressure ratio. Finally, the optimized shapes on the Pareto fronts are investigated, analyzing mechanisms responsible for loss reduction and enhanced compression. Profiles having the lowest losses have flatter camberlines and reduced acceleration of flow on the suction side, while geometries achieving the highest pressure ratio values have a more cambered shape with a concave suction side.

Some Aspects of the Transonic Compressor Tandem Design

For the development of the latest generation of axial compressors, it is necessary to enlarge the design space by using advanced aerodynamic processes. This enables a further increase in efficiency and performance. The use of a tandem blade configuration in a transonic compressor row provides the possibility to enlarge the design space. It is necessary to address the design aspects a bit more in detail in order to efficiently apply this blading concept to turbomachinery. Therefore, in the current study, the known design aspects of tandem blading in compressors will be summed up under the consideration of the aerodynamic effects and construction characteristics of a transonic compressor tandem. Based on this knowledge, a new transonic compressor tandem cascade (DLR TTC) with an inflow Mach number of 0.9 is designed using modern numerical methods and a multi-objective optimization process. Three objective functions as well as three operating points are used in the optimization. Furthermore, both tandem blades and their arrangement are parameterized. From the resulting database of 1246 members, a final best member is chosen as the state-of-the-art design for further detailed investigation. The aim of the ensuing experimental and numerical investigation is to answer the question, whether the tandem cascade resulting from the modern design process fulfills the described design aspects and delivers the requested performance and efficiency criteria. The numerical simulations within the study are carried out with the DLR flow solver TRACE. The experiments are performed at the transonic cascade wind tunnel of DLR in Cologne. The inflow Mach number during the tests is 0.9, and the AVDR is adjusted to 1.3 (design value). Wake measurements with a three-hole probe are carried out in order to determine the cascade performance. The experimental results show an increase in losses and a reduction of the cascade deflection by about 2 deg compared to the design concept. Nevertheless, the experimental and numerical results allow a good understanding of the aerodynamic effects. In addition, planar PIV was applied in a single S1 plane located at midspan to capture the velocity field in the wake of blade 1 in order to analyze the wake flow in detail and describe its influence on the cascade deflection and loss behavior. Finally, an outlook will be given on what future tandem compressor research should be focused.

The Present Challenge of Transonic Compressor Blade Design

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.

Some Aspects of Transonic Compressor Tandem Design

The development of modern axial compressors has already reached a high level. Therefore an enlargement of the design space by means of new or advanced aerodynamic methods is necessary in order to achieve further enhancements of performance and efficiency. The tandem arrangement of profiles in a transonic compressor blade row is such a method. It is necessary to address the design aspects a bit more in detail in order to efficiently apply this blading concept to turbomachinery. Therefore, in the current study the known design aspects of tandem blading in compressors will be summed up under consideration of the aerodynamic effects and construction characteristics of a transonic compressor tandem. Based on this knowledge, a new transonic compressor tandem cascade (DLRTTC) with an inflow Mach number of 0.9 is designed using modern numerical methods and a multi objective optimization process. Three objective functions as well as three operating points are used in the optimization. Furthermore, both tandem blades and their arrangement are parameterized. From the resulting database of 1246 members a final best member is chosen as state-of-the-art design for further detailed investigation. The aim of the ensuing experimental and numerical investigation is to answer the question, whether the tandem cascade resulting from the modern design process fulfills the described design aspects and delivers the requested performance and efficiency criteria. The numerical simulations within the study are carried out with the DLR flow solver TRACE. The experiments are performed at the Transonic Cascade Wind Tunnel of DLR in Cologne. The inflow Mach number during the tests is 0.9 and the AVDR [1, 2] is adjusted to 1.3 (design value). Wake measurements with a 3-hole probe are carried out in order to determine the cascade performance. The experimental results show an increase in losses and a reduction of the cascade deflection by about 2 degrees compared to design concept. Nevertheless, the experimental and numerical results allow a good understanding of the aerodynamic effects. In addition, Planar PIV was applied in a single S1 plane located at midspan to capture the velocity field in the wake of blade 1 in order to analyze the wake flow in detail and describe its influence on the cascade deflection and loss behavior. Finally, an outlook will be given on what future tandem compressor research should be focused.

Effect of Pressure Ratio on Bending Mode Flutter in a Transonic Linear Cascade

Vibration related issues such as flutter have always been a cause of concern for aircraft engine designers. They not only incur unwarranted cost and time overruns, but also significantly compromise performance and can cause structural damage. This phenomenon has become more relevant for the modern aircraft engines, which employ relatively thin, long blade rows to satisfy ever growing demand for a powerful yet compact engine. The tip sections of such blade rows operate with supersonic relative velocity, where prediction of flutter can get challenging due to unsteady flow features like oscillating shocks and their interaction with the blade motion. Linear cascades that represent a specific radial location of the rotor have proven to be a reliable tool for flutter studies. To facilitate flutter experiments at flow Mach numbers realistic to the aircraft engine components, a transonic cascade facility operating at a Mach Number (M) of 1.3 with the ability to oscillate the central blade in the cascade has been developed. The cascade consists of 5 blades and two false blades of which the central blade is oscillated in heave, which represents the bending mode of the rotor. The typical reduced frequencies associated with this kind of flutter in practice (k ∼ 0.1) correspond to a high dimensional frequency of 200 Hz for the present case. A barrel cam mechanism is used to provide such high frequency oscillations. The parameters varied in the present study include the reduced frequency (k) and the static pressure ratio (SPR) across the cascade, which is varied with the help of tailboard and flap arrangement located at the back end of the cascade. Three SPR cases of 1.05, 1.25, and 1.35 are considered and at each of these pressure ratio cases, the reduced frequency is varied. The unsteady loads are measured on the oscillating central blade during the oscillation cycle to quantify the energy transfer from flow to blade and shadowgraphy is used to visualize the shocks. The results from these experiments indicate flutter at lower k values for all the SPR cases tested, while the higher k values are damped. The magnitude of excitation or damping at any particular frequency is also observed to increase with increasing SPR.

Numerical method for an assessment of steady and motion-excited flowfields in a transonic cascade wind tunnel

This article presents a numerical method and its application for an assessment of the flow field inside a wind tunnel. A structured computational fluid dynamics (CFDs) solver with overset mesh technique is developed in order to simulate geometrically complex configurations. Applying the developed solver, a whole transonic cascade wind tunnel is modeled and simulated by a two-dimensional manner. The upstream and downstream periodicity of the cascade and the effect of the tunnel wall on the unsteady flow field are focused on. From the steady flow simulations, the existence of an optimum throttle position for the best periodicity for each tailboard angle is shown, which provides appropriate aerodynamic characteristics of ideal cascades in the wind tunnel environment. Unsteady simulations with blade oscillation is also conducted, and the difference in the influence coefficients between ideal and wind tunnel configurations becomes large when the pressure amplitude increases on the lower blades.