Alexander Hergt

Effects of Vortex Generator Application on the Performance of a Compressor Cascade

The performance of a compressor cascade is considerably influenced by secondary flow effects, like the cross flow on the end wall as well as the corner separation between the wall and the vane. An extensive experimental study of vortex generator application in a highly loaded compressor cascade was performed in order to control these effects and enhance the aerodynamic performance. The results of the study will be used in future projects as a basis for parameterization in the design and optimization process for compressors in order to develop novel nonaxisymmetric endwalls as well as for blade modifications. The study includes the investigation of two vortex generator types with different geometrical forms and their application on several positions in the compressor cascade. The investigation includes a detailed description of the secondary flow effects in the compressor cascade, which is based on numerical and experimental results. This gives the basis for a specific approach of influencing the cascade flow by means of vortex generators. Depending on the vortex generator type and position, there is an impact on the end wall cross flow, the development of the horse shoe vortex at the leading edge of the vane, and the extent of the corner separation achieved by improved mixing within the boundary layer. The experiments were carried out on a compressor cascade at a high-speed test facility at DLR in Berlin at minimum loss (design point) and off-design of the cascade at Reynolds numbers up to Re = 0.6 × 106 (based on 40-mm chord) and Mach numbers up to M = 0.7. At the cascade design point, the total pressure losses could be reduced by up to 9% with the vortex generator configuration, whereas the static pressure rise was nearly unaffected. Furthermore, the cascade deflection could be influenced considerably by vortex generators and also an enhancement of the cascade stall range could be achieved. All these results will be presented and discussed with respect to secondary flow mechanisms. Finally, the general application of vortex generators in axial compressors will be discussed.

Advanced Nonaxisymmetric Endwall Contouring for Axial Compressors by Generating an Aerodynamic Separator—Part I: Principal Cascade Design and Compressor Application

Modern methods for axial compressor design are capable of shaping the blade surfaces in a three-dimensional way. Linking these methods with automated optimization techniques provides a major benefit to the design process. The application of nonaxisymmetric contoured endwalls is considered to be very successful in turbine rotors and vanes. Concerning axial compressors, nonaxisymmetric endwalls are still a field of research. This two-part paper presents the recent development of a novel endwall design. An aerodynamic separator, generated by a nonaxisymmetric endwall groove, interacts with the passage vortex. This major impact on the secondary flow results in a significant loss reduction because of load redistribution, reduction in recirculation areas, and suppressed corner separation. The first paper deals with the development of the initial endwall design using a linear compressor cascade application. A brief introduction of the design methods is provided, including the automated optimization and the 3D process chain with a focus on the endwall contouring tool. Hereafter, the resulting flow phenomena and physics due to the modified endwall surface are described and analyzed in detail. Additionally, the endwall design principal is transferred to an axial compressor stage. The endwall groove is applied to the hub and casing endwalls of the stator, and the initial numerical investigation is presented. For highly loaded operating points, the flow behavior at the hub region can be improved in accord with the cascade results. Obviously, the casing region is dominated by the incoming tip vortex generated by the rotor and still remains an area for further investigations concerning nonaxisymmetric endwall contouring.

Advanced Nonaxisymmetric Endwall Contouring for Axial Compressors by Generating an Aerodynamic Separator—Part II: Experimental and Numerical Cascade Investigation

Modern methods for axial compressor design are capable of shaping the blade surfaces in a three-dimensional way. Linking these methods with automated optimization techniques provides a major benefit to the design process. The application of nonaxisymmetric contoured endwalls is considered to be very successful in turbine rotors and vanes. Concerning axial compressors, nonaxisymmetric endwalls are still a field of research. This two-part paper presents the recent development of a novel endwall design. A vortex created by a nonaxisymmetric endwall groove acts as an aerodynamic separator, preventing the passage vortex from interacting with the suction side boundary layer. This major impact on the secondary flow results in a significant loss reduction by means of load redistribution, reduction in recirculation areas, and suppressed corner separation. Part I of this paper deals with the endwall design and its compressor application. The resulting flow phenomena and physics are described and analyzed in detail. The second paper presents the detailed experimental and numerical investigation of the developed endwall groove. The measurements carried out at the transonic cascade wind tunnel of DLR in Cologne, demonstrated a considerable influence on the cascade performance. A loss reduction and redistribution of the cascade loading were achieved at the aerodynamic design point, as well as near the stall condition of the cascade. This behavior is well predicted by the numerical simulation. The combined analysis of experimental and numerical flow patterns allows a detailed interpretation and description of the resulting flow phenomena. In this context, high fidelity 3D-Reynolds-averaged Navier―Stokes flow simulations are required to analyze the complex blade and endwall boundary layer interaction.

Experimental Investigation of Flow Control in Compressor Cascades

A large part of the total pressure losses in a compressor stage is caused by secondary flow effects like the separation between the wall and the vane i.e., a corner separation. An experimental and numerical investigation in a highly loaded compressor cascade was performed to understand the fluid mechanic mechanism of this corner separation in order to control it by using vortex generators. The experiments were carried out with a compressor cascade at a high-speed test facility at DLR in Berlin. The cascade consisted of five vanes and their profiles represent the cut at 10% of span distance from the hub of the stator vanes of a single stage axial compressor. The experiments were accomplished at Reynolds numbers up to Re = 0.6 × 106 (based on 40 mm chord) and Mach numbers up to M = 0.7. To measure the total pressure losses of the cascade (caused by the corner separation) a wake rake was used. It consisted of 26 pitot probes to measure the total pressure distribution of the outflow and 4 Conrad probes to determine the outflow angles. To detect the separation area on the vane, a flow visualisation technique was used. In addition to the experiments, numerical computations were carried out with the URANS TRACE, which has been developed at DLR for the simulation of steady and unsteady turbomachinery flow. The computations were performed with identical geometrical conditions as in the experiments, including the measured inflow boundary layer conditions at the side walls. The experiments were performed with the aim of controlling the corner separation. In this case, vortex generators as a passive flow control device were used. The vortex generators were attached at the surface of the suction side of the vanes. The flow control device is producing a strong vortex, which enhances the mixing between the main flow and the retarded boundary layer at the side wall. Thus, the corner separation is reduced on the vanes. The experiments were carried out at the peak efficiency (design point) of the cascade in order to optimize the design of the vortex generators for an application in turbomachines.Copyright

Loss Reduction in Compressor Cascades by Means of Passive Flow Control

Secondary flow effects like the corner stall between the wall and the vane in a compressor stage are responsible for a large part of total pressure losses. An extensive experimental study of flow control in a highly loaded compressor cascade was performed in order to decrease the separation and reduce the losses by means of vortex generators. The vortex generators were attached at the surface of the cascade side walls. These flow control devices produce strong vortices, which enhance the mixing between the main flow and the decelerated boundary layer at the side wall. Thus, the corner flow separation and the total pressure losses could be reduced. The experiments were carried out with a compressor cascade at a high-speed test facility at the DLR in Berlin at minimum loss (design point) and off-design of the cascade at Reynolds numbers up to Re = 0.6 × 106 (based on 40 mm chord) and Mach numbers up to M = 0.7. The cascade consisted of five vanes. The blade profiles are comparable to the hub section of the stator vanes used in the transonic compressor test rig running at Technische Universitat Darmstadt. In the range between −2° and +4° angle of incidence the total pressure losses of the cascade could be reduced up to 4.6% by means of vortex generators, whereas the static pressure rise was not influenced. Based on the results of the cascade measurements, the vortex generators were applied in front of the stator row of the single stage axial compressor at Technische Universitat Darmstadt. A numerical simulation of the compressor flow provided an indication for the adjustment of the vortex generators at the hub and casing. In the experiments the pressure rise and the efficiency of the axial compressor was measured and it could be shown that vortex generators partially improve the efficiency.Copyright

Analyzing and optimizing geometrically degraded transonic fan blades by means of 2D and 3D simulations and cascade measurements

The present study deals with the influence of geometrically degraded transonic engine fan blades on the fan’s aerodynamic behavior. The study is composed of three phases; the first consists of 3D simulations to point out changes in the performance parameters caused by blade degradations. In the second phase, 2D optimizations are carried out to determine the potential of redesigning the blade and in the third phase, measurements on a transonic cascade are used to experimentally verify the numeric results.During engine operation as well as maintenance processes, geometric variations of the fan blades, and especially of the blades’ leading edges, are observed. They mainly originate from the ambient conditions under which the engine is operated. Though the deformations of the blade differ widely, several typical degradation types can be identified. In advance of the study, these degradation types have been systematized and simplified models representing different degrees of degradation have been built.In the first phase, the models are aerodynamically analyzed by means of 3D simulations. A high influence on the performance parameters is found for a fan blade exposed to long-term erosion. The model’s characteristics are a blunt leading edge and a reduced chord length. In contrast, the performance parameters of a model representing a re-contoured blade (reduced chord length but reshaped leading edge) are shown to be similar to those of a new fan blade. This leads to the conclusion that an eroded blade may offer almost the initial performance parameters as long as the leading edge is well reshaped.Since the model of the long-term eroded blade shows great changes in the fan’s performance and the best optimization potential, this has been chosen for the further analysis in the following phases.In the second phase, 2D optimizations are applied to three airfoil sections at different heights of the blade. The parameterization used is limited to a small area of the leading edge; the shape of the rest of the blade is kept constant. The optimizations lead to loss reduction and demonstrate the potential of the optimization process.The third phase is carried out in the Transonic Cascade Wind Tunnel of the Institute of Propulsion Technology in Cologne. As the transonic part of the fan blade is the most sensitive to geometric changes, a transonic airfoil with long-term erosion has been chosen. During the tests, the following measurement techniques are applied: Static pressure probes to determine the Mach number distribution, a 3-hole probe to detect exit angle and loss distribution, Schlieren photographs and PIV-measurements to locate the shock system, the L2F method to measure the cascade inflow angle and to resolve the boundary layer distribution and Liquid crystal measurements to observe transition activities. The full analysis of the measurements with PIV, L2F and Liquid Crystals are still in progress, but the evaluation of the loss polar and the Schlieren photographs show increased losses for the degraded blade and a good match with the numeric results.Copyright

A NEW APPROACH FOR COMPRESSOR ENDWALL CONTOURING

Against the background of the high development status of modern axial compressors, a further performance enhancement is linked with the extension of the design space in the development process and the concentration on the essential loss mechanisms in the compressor. The performance of a compressor cascade is considerably influenced by secondary flow effects in the near endwall region, since up to 50 percent (for low aspect ratio) of the losses in the bladed channel of a turbomachinery are linked to the endwalls. In this context the application of non-axisymmetric profiled endwalls provides a potential for compressor improvement. The paper presents the detailed experimental and numerical investigation of controlling the endwall cross flow in a compressor cascade. The general approach is based on a boundary layer fence arrangement, whose application on the compressor endwall works as a non-axisymmetric endwall contour. This non-axisymmetric endwall modification constrains the interaction of the endwall cross flow with the suction side boundary layer, thus the onset of the corner separation is delayed and a significant loss reduction of 8 percent is achieved. The experiments were carried out in a linear compressor cascade at the high-speed cascade wind tunnel of the DLR in Berlin at peak efficiency (design point) and off-design of the cascade at Mach number M = 0.67. Furthermore, high fidelity 3D-RANS flow simulations were performed in order to analyze the complex blade and endwall boundary layer interaction. The combined consideration of experimental and numerical flow pattern allows a detailed interpretation and description of the resulting flow phenomena.Copyright

About Transonic Compressor Tandem Design: A Principle Study

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. For an efficient industrial application the knowledge of the fundamental design principles is needed. This paper presents the recent research work on transonic compressor tandem profiles at DLR Institute of Propulsion Technology. It deals with the fundamental description of the operation principles of a modern transonic compressor tandem cascade. By considering these principles and based on an optimization database with over 1200 members design recommendations are developed

Endwall Contouring and Fillet Design for Reducing Losses and Homogenizing the Outflow of a Compressor Cascade

The following paper deals with the development of an optimized fillet and an endwall contour for reducing the total pressure loss and for homogenizing the outflow of a highly loaded cascade with a low aspect ratio. The NACA-65 K48 cascade profile without a fillet and without endwall contouring is used as a basis. Optimizations are performed using the DLR in-house tool AutoOpti and the RANS-solver TRACE. Three operating points at an inflow Mach number of 0.67 with different inflow angles are used to secure a wide operating range of the optimized design. At first only a fillet is optimized. The optimized fillet is small at the leading edge and rather high, wide and thick towards the trailing edge. It reduces the total pressure loss and homogenizes the outflow up to a blade height of 20 %. Following this a combined optimization of the endwall and the fillet is performed. The optimized contour leads to the development of a vortex, which changes the secondary flow in such a way, that the corner separation is reduced, which in turn significantly reduces the total pressure loss up to 16 % in the design operating point. The contour in the outflow region leads to a significant homogenization of the outflow in the near wall region.

Advanced Non-Axisymmetric Endwall Contouring for Axial Compressors by Generating an Aerodynamic Separator—Part I: Principal Cascade Design and Compressor Application

Modern methods for axial compressor design are capable of shaping the blade surfaces in a three dimensional way. Linking these methods with automated optimization techniques provides a major benefit to the design process. The application of non-axisymmetric contoured endwalls is considered to be very successful in turbine rotors and vanes. Concerning axial compressors non-axisymmetric endwalls are still a field of research. This two-part paper presents the recent development of a novel endwall design. An aerodynamic separator, generated by a non-axisymmetric endwall groove, interacts with the passage vortex. This major impact on the secondary flow results in a significant loss reduction because of load redistribution, reduction of recirculation areas and suppressed corner separation. The first paper deals with the development of the initial endwall design using a linear compressor cascade application. A brief introduction of the design methods is provided, including the automated optimization, the 3D process chain with a focus on the endwall contouring tool. Hereafter the resulting flow phenomena and physics due to the modified endwall surface are described and analyzed in detail. Additionally, the endwall design principal is transferred to an axial compressor stage. The endwall groove is applied to the hub and casing endwalls of the stator and the initial numerical investigation is presented. For highly loaded operating points the flow behaviour at the hub region can be improved in accord with the cascade results. Obviously, the casing region is dominated by the incoming tip vortex generated by the rotor and still remains an area for further investigations concerning non-axisymmetric endwall contouring.Copyright