Different active flow control techniques have been investigated in a 1.5-stage axial-flow compressor. Looking at a low-speed single-stage environment, many researchers have shown that highly loaded compressors are tip critical, showing stall inception caused by short length scale disturbances (spikes). It has been shown by several authors that these disturbances are related to the spillage of endwall flow ahead of the blading (spill forward). For the present work, different tip injection configurations were investigated in order to stabilize the near casing flow, increasing the operating range of the compressor. Stall margin improvement and the impact on stage efficiency are compared and discussed. Oil flow pictures of the casing wall above the rotor and of the stator blades as well as traverse data from pneumatic 5-hole probes show the impact of flow control on rotor and stator performance. Another method of energizing the casing wall boundary layer is the removal of low energy fluid by a circumferential slot above the rotor, which was also studied experimentally. Again, the impact on compressor operating range and efficiency, as well as flow field information collected by oil flow visualization and traverse data are discussed. Comparing the different flow control techniques, it is shown that increasing stall margin is not directly linked to stage efficiency. As described in various publications, discrete tip injection is a very powerful technique as far as range extension is concerned, but it also has substantial drawbacks such as the circumferential inhomogeneity of the rotor exit flow. These inhomogeneities may result in poor stator performance, overall resulting in a drop of stage efficiency. This problem does not occur if circumferential boundary layer removal above the rotor is used. This method however shows much less potential for increasing the operating range.
The following paper describes an experimental investigation of a highly loaded stator cascade with a pitch to chord ratio of t/l=0.6. Experiments without as well as with active flow control by means of endwall and suction side blowing were conducted. Five-hole-probe measurements in pitchwise and spanwise directions as well as endwall oil flow visualizations were carried out in order to determine the performance of the cascade and to analyze the flow phenomena occurring. To quantify the effectivity of the active flow control method, taking the additional energy input into account, corrected losses and an efficiency, which relates the difference of flow power deficit with and without active flow control to the flow power of the blowing jet itself, were evaluated. Even though an increase of static pressure rise could be achieved, a decrease of the total pressure losses was possible for a few operating points only.
The following paper describes an experimental investigation of a highly loaded stator cascade with a pitch to chord ratio of t/l = 0.6. Experiments without as well as with active flow control by means of endwall and suction side blowing were conducted. Five-hole-probe measurements in pitchwise and spanwise direction as well as endwall oil flow visualizations were carried out in order to determine the performance of the cascade and to analyze the flow phenomena occuring. To quantify the effectivity of the active flow control method, taking the additional energy input into account, corrected losses and an efficiency, which relates the difference of flow power deficit with and without active flow control to the flow power of the blowing jet itself, were evaluated. Even though an increase of static pressure rise could be achieved, a decrease of the total pressure losses was possible for a few operating points only.
This paper provides a method to transfer geometric uncertainties of compressor blades into the numerical simulation. Therefore a method to capture geometric variations of measured blades by typical profile parameters is introduced. An optical measurement technique using structured light is applied to scan compressor blades in order to receive a three–dimensional point cloud of the measured blade. The evaluation of these points is done on curves of constant spanwise coordinate between hub and casing. In this way, section outlines are extracted, which then are split into camber lines and chord lines. The derived thickness and camber distributions are used to specify typical profile parameters for each section. To consider the geometric uncertainties in numerical simulation, the design geometry is adapted through a special reconstruction algorithm. Therefore the differences between the measured airfoil and the design geometry are quantified by the profile parameters. Since only the difference is analyzed, few parameters are needed to model the measured geometry. The three-dimensional blade then is reconstructed through assignment of the parameters to the spanwise coordinate. To illustrate the developed method, the whole process chain is applied on a selected compressor blade.
A reduction of CO2 emissions of aero engines can only be achieved by a reduction in the fuel consumption. For turbofan engines a major key for this is an increase in the engine bypass ratio to enhance the propulsive efficiency. This leads to an increased mass flow in the bypass duct and hence higher contributions of the bypass duct system to the flow pressure losses and hence the engine fuel consumption. Additionally the turbofan core engine gets smaller in size and the overall diameter of the engine increases. This requires advanced engine mounting concepts, such as core mounted subsystems, which influence the arrangement and the mechanical and aerodynamic design of components in the bypass duct such as fan outlet guide vane, struts, fairings or bifurcations. This requires an optimized design of the turbofan bypass system with regards to structural, aerodynamic and acoustic criteria. The topology of the bypass duct as well as the position and the design of the individual components have a significant impact on the weight, the efficiency and the noise radiation of the engine and hence need to be investigated in a systematic approach. Such a design approach was developed in the R&T project OPAL, led by Rolls-Royce Deutschland and funded by the German Federal State of Brandenburg. The design approach covers the following steps: - Optimization of the bypass duct shape with regards to minimum pressure loss taking into consideration the mechanical requirements of an engine with core mounted subsystems. - Optimization of strut, fairing and bifurcation shapes with regards to minimum pressure loss, robustness against flow deviations and minimized upstream flow effects taking into consideration structural and mechanical requirements by engine loads and subsystem routings. - Optimization of fan outlet guide vane profile and plan form shape with regards to minimum pressure losses and maximum working range taking into consideration structural and acoustic constraints. - Optimization of fan-outlet guide vane, strut and bifurcation interaction with regards to minimum pressure losses and maximized aerodynamic stability of the fan and outlet guide vanes. The current paper will present the design approach, the optimization processes and the results of the optimization of the turbofan bypass duct system for the application on modern high-bypass ratio aero engines.
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