No abstract

The stable operating range of a centrifugal compressor stage of an engine turbocharger is limited at low mass flow rates by aerodynamic instabilities which can lead to the onset of rotating stall or surge. There have been many techniques employed to increase the stable operating range of centrifugal compressor stages. The literature demonstrates that there are various possibilities for adding special treatments to the nominal diffuser vane geometry, or including injection or bleed flows to modify the diffuser flow field in order to influence diffuser stability. One such treatment is the porous throat diffuser (PTD). Although the benefits of this technique have been proven in the existing literature, a comprehensive understanding of how this technique operates is not yet available. This paper uses experimental measurements from a high pressure ratio (PR) compressor stage to acquire a sound understanding of the flow features within the vaned diffuser which affect the stability of the overall compression system and investigate the stabilizing mechanism of the porous throat diffuser. The nonuniform circumferential pressure imposed by the asymmetric volute is experimentally and numerically examined to understand if this provides a preferential location for stall inception in the diffuser. The following hypothesis is confirmed: linking of the diffuser throats via the side cavity equalizes the diffuser throat pressure, thus creating a more homogeneous circumferential pressure distribution, which delays stall inception to lower flow rates. The results of the porous throat diffuser configuration are compared to a standard vaned diffuser compressor stage in terms of overall compressor performance parameters, circumferential pressure nonuniformity at various locations through the compressor stage and diffuser subcomponent analysis. The diffuser inlet region was found to be the element most influenced by the porous throat diffuser, and the stability limit is mainly governed by this element.

Vaned diffusers in centrifugal compressor stages are used to achieve higher stage pressure ratios, higher stage efficiencies, and more compact designs. The interaction of the stationary diffuser with the impeller can lead to resonant vibration with potentially devastating effects. This paper presents unsteady diffuser vane surface pressure measurements using in-house developed, flush mounted, fast response piezoresistive pressure transducers. The unsteady pressures were recorded for nine operating conditions, covering a wide range of the compressor map. Experimental work was complemented by 3D unsteady computational fluid dynamics (CFD) simulations using ansys cfx V12.1 to detail the unsteady diffuser aerodynamics. Pressure fluctuations of up to 34.4% of the inlet pressure were found. High pressure variations are present all along the vane and are not restricted to the leading edge region. Frequency analysis of the measured vane surface pressures show that reduced impeller loading, and the corresponding reduction of tip leakage fluid changes the characteristics of the fluctuations from a main blade count to a total blade count. The unsteady pressure fluctuations in the diffuser originate from three distinct locations. The impact of the jet-wake flow leaving the impeller results in high variation close to the leading edge. It was observed that CFD results overpredicted the amplitude of the pressure fluctuation on average by 62%.

The main centrifugal compressor performance criteria are pressure ratio, efficiency, and wide flow range. The relative importance of these criteria, and therefore the optimum design balance, varies between different applications. Vaned diffusers are generally used for high-performance applications as they can achieve higher efficiencies and pressure ratios, but have a reduced operating range, in comparison to vaneless diffusers. Many impeller-based casing treatments have been developed to enlarge the operating range of centrifugal compressors over the last decades but there is much less information available in open literature for diffuser focused methods, and they are not widely adopted in commercial compressor stages. The development of aerodynamic instabilities at low mass flow rate operating conditions can lead to the onset of rotating stall or surge, limiting the stable operating range of the centrifugal compressor stage. More understanding of these aerodynamic instabilities has been established in recent years. Based on this additional knowledge, new casing treatments can be developed to prevent or suppress the development of these instabilities, thus increasing the compressor stability at low mass flow rates. This paper presents a novel vaned diffuser casing treatment that successfully increased the stable operating range at low mass flow rates and high pressure ratios. Detailed experimental measurements from a high pressure ratio turbocharger compressor stage combined with complementary CFD simulations were used to examine the effect of the new diffuser casing treatment on the compressor flow field and led to the improvement in overall compressor stability. A detailed description of how the new casing treatment operates is presented within the paper.

The present numerical study aims at examining the influence of intrarow interaction effects in aerodynamic damping predictions of an axial turbine rotor. The investigated operating point corresponds to a resonance crossing associated with the fundamental engine order of the stator blade row. Accordingly, the pressure perturbations induced by the vibration of the rotor at its modal frequency are found to be coincident in frequency and thus superimpose with the pressure perturbations resulting from intrarow interaction phenomena. A methodology for extracting vibration induced pressure perturbations for the subsequent calculation of the vibration induced modal aerodynamic damping is established and applied within the scope of the present study. Applying this methodology, both the influence of the underlying mean and transient flow field as well as the influence of acoustic wave reflections at the adjacent stator blade row is investigated on the predicted aerodynamic damping. In this context, the underlying mean flow field, which is found to be slightly altered in the presence of intrarow interaction phenomena, was proven to have a significant influence on vibration induced pressure perturbations. Moreover, acoustic wave reflections at the adjacent stator blade row are found to have the capability of influencing the aerodynamic damping depending on their actual phasing when impinging onto the turbine rotor.