International Journal of Turbomachinery, Propulsion and Power — Open Access Journal

The LS89 high pressure axial turbine vane was originally designed and optimized for a downstream isentropic Mach number of 0.9. This profile has been widely used for computational fluid dynamics (CFD) validation in the open literature but very few attempts have been made to improve the already optimized design. This paper presents a sound methodology to design and optimize the LS89 using computer-aided design (CAD) at design conditions. The novelty of the study resides in the parametrization of design space, which is done at the CAD level, and the detailed analysis of the aerodynamic performance of the optimized design. Higher level constraints are imposed on the shape, such as the trailing edge thickness, the axial chord length, and G2 geometric continuity between the suction side and pressure side at the leading edge. The gradients used for the optimization are obtained by applying algorithmic differentiation to the CAD kernel and grid generator and the discrete adjoint method to the CFD solver. A reduction of almost 12% entropy generation is achieved, which is equivalent to a 16% total pressure loss reduction. The entropy generation is reduced while keeping the exit flow angle as a flow constraint, which is enforced via the penalty formulation. The resulting unconstrained optimization problem is solved by the L-BFGS-B algorithm. The flow is governed by the Reynolds-averaged Navier-Stokes equations and the one-equation transport Spalart-Allmaras turbulence model. The optimal profile is compared and benchmarked against the baseline case. Full article

For the numerical prediction of turbomachinery flows, a two-equation turbulence model in combination with a proper transition model to account for laminar boundary layers and their transition to turbulence is state of the art. This paper presents the ability of such a method (k-ω + γ-Re_Θ) for turbulence prediction and the effect on three-dimensional boundary layer behavior. For this purpose, both applied models (turbulence and transition) are improved to better account for turbulence length scale effects and three-dimensional transition prediction (Bode et al., 2014 and 2016), since these are the main deficiencies in predicting such kinds of flows. The improved numerical method is validated and tested on existing turbine cascades with detailed experimental data for the viscous regions and additionally on a low-speed axial compressor rig where wake-induced transition takes place. Full article

During the shutdown of a gas turbine, natural convection flows occur inside of the engine casing. These natural convection flows cause the top of the engine casing to cool more slowly than the bottom of the casing. This non-uniform cooling can cause the casing to distort, known as “cat-back” distortion. This distortion leads to reduced tip and seal clearances, which can cause rubbing and binding. In this study, two different methods for reducing the non-uniform casing heat transfer in an industrial gas turbine during shutdown were investigated. A baseline case with no active method for reducing the casing heat transfer was compared with purge and flow extraction methods. It was found that the new proposed extraction method improved the uniformity of the casing heat transfer during the initial shutdown period. Full article

State-of-the-art gas turbines (GT) operate at high temperatures that exceed the endurance limit of the material, and therefore the turbine components are cooled by the air taken from the compressor. The cooling provides a positive impact on the lifetime of GT but has a negative impact on its performance. In convection-cooled turbine blades the coolant is usually discharged through the trailing edge and leads to limitations on the minimal size of the trailing edge, thereby negatively affecting the losses. Moreover, the injection of cooling air in the turbine disturbs the main flow, and may lead to an additional increase in loss. Trailing edge loss is a significant part of the overall loss in modern gas turbines. This study comprises investigations of the unguided flow angle, the trailing edge shape, and cooling air injection through the trailing edge on the base pressure and profile losses in cooled blades. Some vane and blade cascades with different unguided turning angle and two shapes of trailing edges with and without coolant injection were studied both experimentally and numerically. This analysis provides a split of losses caused by different factors, and offers opportunities for efficiency and lifetime improvements of real engine designs/upgrades. In particular, it is shown that an increase in the unguided turning angle and the use of a round trailing edge result in a reduction of loss in case of a relatively thick trailing edge. Numerical investigation showed that an increase in the unguided turning angle at the initial transonic vane with a thick and blunt trailing edge, without a change in other basic geometric parameters, allowed for a significant reduction of the profile loss by about 3–4% at the exit Mach number M_2is = 0.7–1.0. Experimental investigation of four cascades with cooling air injection into the base flow through the trailing edge allowed us to validate the fact that in blades with a low level of base pressure C_pb < −0.1 at $\overline{m}$ = 0 a non-monotonic dependence of the change of losses against relative cooling air mass flow $\overline{m}$ is observed. Firstly, the cooling air injection into wake increases base pressure and decreases losses; then the losses start to increase with increasing cooling mass flow due to the interaction between the main flow and the cooling air (mixing losses) and, finally, due to the cooling mass flow increase and momentum increase losses are decreased. In blades with an increased level of the base pressure coefficient C_pb ≥ −0.1 at $\overline{m}$ = 0 the cooling air injection results in an increase in losses right from the beginning of the injection and then, according to the cooling mass flow increase and momentum rise, losses decrease. It is also shown that injection through the trailing edge slot parallel to the main flow leads to a neutral loss impact and even a loss reduction in the subsonic range and a loss increase in the supersonic range of exit Mach numbers. Full article

Current civil aviation is characterized by rising cost and competitive pressure, which is partly passed to the MRO (Maintenance, Repair and Overhaul) companies. To improve the maintenance, condition-based maintenance is established, which is characterized by tailored maintenance actions for each part of the jet engine, depending on the individual engine history and operating conditions. Thereby, prediction models help engineers to authorize maintenance actions as effectively as possible. This paper will help to improve these prediction models. Therefore, the influence of specific deterioration of a high pressure compressor (HPC) to jet engine performance parameters such as exhaust gas temperature (EGT) and specific fuel consumption (SFC) will be investigated. For this purpose, computational fluid dynamic (CFD) calculations of deteriorated HPC geometries are carried out and serve as a basis to scale the reference HPC performance characteristics to deteriorated ones. To evaluate the changes in performance parameters, a modular performance synthesis model is set up. In this model, the HPC map is exchanged with deteriorated ones. As a result, the influence of geometric deviations to the design intent can be determined, and the MRO companies are able to focus on the most relevant sections of the compressor blading. Full article

The decreasing performance of jet engines during operation is a major concern for airlines and maintenance companies. Among other effects, the erosion of high-pressure compressor (HPC) blades is a critical one and leads to a changed aerodynamic behavior, and therefore to a change in performance. The maintenance of BLISKs (blade-integrated-disks) is especially challenging because the blade arrangement cannot be changed and individual blades cannot be replaced. Thus, coupled deteriorated blades have a complex aerodynamic behavior which can have a stronger influence on compressor performance than a conventional HPC. To ensure effective maintenance for BLISKs, the impact of coupled misshaped blades are the key factor. The present study addresses these effects on the aerodynamic performance of a first-stage BLISK of a high-pressure compressor. Therefore, a design of experiments (DoE) is done to identify the geometric properties which lead to a reduction in performance. It is shown that the effect of coupled variances is dependent on the operating point. Based on the DoE analysis, the thickness-related parameters, the stagger angle, and the max. profile camber as coupled parameters are identified as the most important parameters for all operating points. Full article

The paper presents an experimental campaign aimed at the characterization of the relationship between cavitation-induced instabilities and forces acting on the shaft, relevant to space application turbopumps. The experiments have been carried out on a six-bladed unshrouded centrifugal turbopump. Pressure fluctuations are analyzed in their frequency content for understanding the instability nature (axial, rotating) and their main characteristics (e.g., amplitude, rotating direction). The spectral analysis of the force components highlights a strong relationship of the z-component (along the rotating axis) with axial instabilities. On the other hand, rotating cavitation may involve force oscillations along all the three components leading to unwanted and dangerous fluctuating unbalances perpendicular to the rotating axis. Full article

Efficient aero engine operation requires not only optimized components like compressor, combustor, and turbine, but also an optimal balance between these components. Therefore, a holistic coupled optimization of the whole engine involving all relevant components would be advisable. Due to its high complexity and wide variety of design parameters, however, such an approach is not feasible, which is why today’s aero engine design process is typically split into different component-specific optimization sub-processes. To guarantee the final functionality, components are coupled by fixed aerodynamic and thermodynamic interface parameters predefined by simplified performance calculations early in the design process and held constant for all further design steps. In order not to miss the optimization potential of variable interface parameters and the unlimited design space on higher-fidelity design levels, different coupling and optimization strategies are investigated and demonstrated for a reduced compressor-combustor test case problem by use of 1D and 2D aero design tools. The new holistic design approach enables an exchange of information between components on a higher-fidelity design level than just simple thermodynamic equations, as well as the persecution of global engine design objectives like efficiency or emissions, and provides better results than separated component design with fixed interfaces. Full article

This paper models the kinematics of the vortex system of an encased axial turbomachine at part load and overload applying analytical methods. Thus far, the influence of the casing and the tip clearance on the kinematics have been solved separately. The vortex system is composed of a hub, bound and tip vortices. For the nominal operating point $\phi \approx {\phi }_{\mathrm{opt}}$ and negligible induction, the tip vortices transform into a screw. For part load operation $\phi \to 0$ the tip vortices wind up to a vortex ring, i.e., the pitch of the screw vanishes. The vortex ring itself is generated by bound vortices rotating at the angular frequency $\Omega$ . The hub vortex induces a velocity on the vortex ring causing a rotation at the sub-synchronous frequency ${\Omega }_{\mathrm{ind}}=0.5\Omega$ . Besides, the vortex ring itself induces an axial velocity. Superimposed with the axial main flow this results in a stagnation point at the tube wall. This stagnation point may wrongly be interpreted as dynamic induced wall stall. For overload operation $\phi \to \infty$ the vortex system of the turbomachine forms a horseshoe, i.e., the pitch of the screw becomes infinite. Both hub and tip vortices are semi-infinite, straight vortex filaments. The tip vortices rotate against the rotating direction of the turbomachine due to the induction of the hub vortex yielding the induced frequency ${\Omega }_{\mathrm{ind}}=-0.5\Omega /s$ with the tip clearance s. Full article

Large eddy simulations were carried out in order to investigate the influence of unsteady incoming wakes with different profiles on the loss mechanisms of the high lift T106Alinear low-pressure turbine (LPT) cascade. Bars placed upstream of the LPT blade were set into rotation around their axis, thus generating circulation, as well as asymmetrical wake profiles. Three different rotation rates were simulated, yielding different wake parameters that were then compared to an actual turbine blade wake profile. Whereas the commonly-used non-rotating bars generated wakes with turbulent kinetic energy levels several times higher than that of an actual blade wake, the case with counter-clockwise rotation led to more rapid wake mixing. All three wakes were able to trigger boundary layer transition and thus intermittently prevent separation on the suction surface. However, the weaker the wakes, the larger and longer lasting the separation bubbles became, and an increase in profile losses could be observed. Interestingly, the configuration with the weakest wake and the largest separation bubble resulted in a reduction of the overall LPT loss. Full article

In radial pumps and turbines, the centrifugal through-flow in both the front and the back chambers is quite common. It strongly impacts the core swirl ratio, pressure distribution, axial thrust and frictional torque. In order to investigate these relationships experimentally, a test rig was designed at the University of Duisburg-Essen and described in this paper. Based on both the experimental and numerical results, correlations are determined to predict the impacts of the centrifugal through-flow on the core swirl ratio, the thrust coefficient and the moment coefficient. Two correlations respectively are determined to associate the core swirl ratio with the local through-flow coefficient for both Batchelor type flow and Stewartson type flow. The correlations describing the thrust coefficient and the moment coefficient in a rotor-stator cavity with centripetal through-flow (Hu et al., 2017) are modified for the case of centrifugal through-flow. The Daily and Nece diagram distinguishing between different flow regimes in rotor-stator cavities is extended with a through-flow coordinate into 3D. The achieved results provide a comprehensive data base which is intended to support the calculation of axial thrust and moment coefficients during the design process of radial pumps and turbines in a more accurate manner. Full article

A highly loaded turbine exit guide vane with active boundary layer control was investigated experimentally in the High Speed Cascade Wind Tunnel at the University of the German Federal Armed Forces, Munich. The experiments include profile Mach number distributions, wake traverse measurements as well as boundary layer investigations with a flattened Pitot probe. Active boundary layer control by fluidic oscillators was applied to achieve improved performance in the low Reynolds number regime. Low solidity, which can be applied to reduce the number of blades, increases the risk of flow separation resulting in increased total pressure losses. Active boundary layer control is supposed to overcome these negative effects. The experiments show that active boundary layer control by fluidic oscillators is an appropriate way to suppress massive open separation bubbles in the low Reynolds number regime. Full article

This article reports on systematic numerical studies on an axial compressor stage with distorted inflow. Four operating points at two speedlines were simulated with an inflow distortion generated by a ${120}^{\circ }$-sector segment with a wedge-type cross-section. With this setup, the interaction between the distorted inflow and the rotor flow was studied. The focus was put on the differences of the interaction between the distorted inflow and rotor flow as well as on the compressor behaviour at subsonic and transonic blade tip speeds, as the general mechanisms have been analysed in more detail in previous publications. The distorted flow itself is not influenced by the blade tip speed, but the interaction phenomena depend strongly on the spool speed and operating point. Additionally, the blade tip speed influences the circumferential sector of the compressor stage exit, which is affected by the distorted flow. The impact reaches from a small sector at 65% spool speed, with the peak efficiency operating point up to nearly the entire annulus at 100% spool speed, near the stall operating point. Full article

Flow measurement using a linear compressor or turbine cascade is a well-established technique to characterize the flow in turbomachines with a certain degree of abstraction. A common way to obtain a general characterization of the flow is to measure the flow downstream of the cascade with a five-hole probe, obtaining, e.g., total pressure losses and flow turning. Pneumatic five-hole probes are used to capture steady or time-averaged flow quantities, if not specified otherwise. In dependency of probe geometry, measurement set-up and flow properties, such measurements can be very time-consuming. Various techniques, in order to decrease the measurement time, are proposed in literature but for certain applications the efforts required to implement such techniques can outweigh the enhanced measurement speed. In this paper, methods proposed by other authors are combined and extended to allow for fast or transient five-hole probe measurements at strongly varying flow conditions. The effectiveness of this method is presented for flow measurements downstream of a compressor cascade with attached and stalled flow (by varying the Reynolds number) as well as with steady and periodically unsteady inflow. The new method allows to reduce the measurement time by up to 90 percent without compromising measurement accuracy. In fact, due to higher spatial resolution, the flow downstream of the cascade can be better resolved with the new method. Full article

Secondary flow and loss development in the T106Div-EIZ low-pressure turbine cascade are investigated utilizing (U)RANS simulations in cases with and without periodically incoming wakes at $M{a}_{2th}=0.59$ and $R{e}_{2th}=2\times {10}^5$. The predictions are compared to experimental data presented by Kirik and Niehuis (2015). The axial mid-span and overall loss development in the T106Div-EIZ and the T106A-EIZ in the steady case are analyzed regarding the effects caused by the different loading distributions and by the divergent endwall geometry. Furthermore, the entropy generation is analyzed in the T106Div-EIZ with periodically incoming wakes in several axial positions of interest and compared to the undisturbed steady case. It is found that in the front-loaded T106Div-EIZ, the incoming wakes cause a premature endwall loss production in the front part of the passage, resulting in a lower intensity of the secondary flow downstream of the passage and a redistribution of the loss generation components. Full article

The continuing rise in turbine entry temperatures has necessitated the development of ever-more advanced cooling techniques. Effusion cooling is an example of such a system and is characterised by a high density of film cooling holes that operate at low blowing ratios, thereby achieving high overall cooling effectiveness. This paper presents both an experimental and computational investigation into the cooling performance of effusion systems. Two flat-plate geometries (with primary hole pitches of 3.0D and 5.75D) are experimentally investigated via a pressure sensitive paint technique yielding high resolution film effectiveness distributions via heat-mass transfer analogy. A computational fluid dynamics (CFD) scalar tracking method was used to model the setup computationally with the results comparing favourably to those obtained from the experiments. The CFD domain was modified to assess the cooling performance from a single film hole ejection. A superposition method was developed and applied to the resulting two-dimensional film effectiveness distribution that quickly yielded data for an array of closely-packed holes, allowing a rapid assessment of a multi-hole effusion type setup. The method produced satisfactory results at higher pitches, but at lower pitches, high levels of jet interactions reduced the performance of the superposition method. Full article

The decrease in the performance of centrifugal compressors operating at low Reynolds numbers (e.g., unmanned aerial vehicles at high altitudes or small turbomachines) can reach 10% due to increased friction. The purposes of this review are to represent the state-of-the-art of the active and passive flow control methods used to improve performance and/or widen the operating range in numerous engineering applications, and to investigate their applicability in low-Reynolds-number centrifugal compressors. The applicable method should increase performance by reducing drag, increasing blade loading, or reducing tip leakage. Based on the aerodynamic and structural demands, passive methods like riblets, squealers, winglets and grooves could be beneficial; however, the drawbacks of these approaches are that their performance depends on the operating conditions and the effect might be negative at higher Reynolds numbers. The flow control method, which would reduce the boundary layer thickness and reduce wake, could have a beneficial impact on the performance of a low-Reynolds-number compressor in the entire operating range, but none of the methods represented in this review fully fulfil this objective. Full article