Abstract Conventional Brayton cycles have demonstrated to be significantly less efficient than alternative propulsion systems (spark ignition, diesel, fuel cells, etc.) for low power output applications, such as for small size UAVs. The gas turbine performance could be enhanced through the introduction of heat exchangers, with the consequent increase of the overall engine weight. Semi-closed cycles have documented advantages of higher thermal efficiency and degree of compactness than traditional intercooled-recuperated open cycles. This paper discusses advantages and applicability of semi-closed cycles to a small gas turbine, designed for a medium altitude UAV mission. In particular, size and altitude effects have been accounted in the performance evaluation of two different semi-closed cycle arrangements designed for an output shaft power of 100 hp (74.57 kW). Resultant performance has been compared with equivalent simple recuperated and intercooled-recuperated open cycles. Furthermore, a final engine performance comparison has been made with data obtained from a similar analysis performed on a larger engine, with a power output of 300 hp (223.71 kW) and designed for an extremely high altitude UAV application. While promising results have been obtained for the larger case study, where semi-closed cycles have demonstrated superior performance and higher engine compactness than conventional solutions, similar trends have not been displayed for the smaller engines, as consequence of the strong size effects observed in the turbomachinery performance. For the 100 hp engine the semi-closed cycles are slightly outperformed by the open cycle engines.

Abstract Rake probes are commonly used in turbomachinery applications to measure distorted inlet flows, including both pressure and swirl distortions. An especially common configuration is the 40 probe rake array, consisting of eight identical arms equally spaced circumferentially around the inlet, each arm having five sensing locations spaced along centers of equal area. By measuring many locations simultaneously, rake probes have the advantage of expedited data collection when compared to traversing probes. However, this reduction in test time comes at the cost of rigid geometry with limited measurement locations. As a result, it is possible for rake probes to miss or exaggerate significant areas of the flow profile, such as large gradients or small features, based on the fixed location of the probe and the particular details of the distortion. The purpose of this paper is to demonstrate a procedure that can be used to evaluate the ability of any desired probe configuration (40 probe rake or otherwise) to sufficiently and accurately measure a non-uniform flow profile. Results of this procedure for a range of profiles and probe configurations are also presented. In order to accurately determine the impact of discrete sampling on the results, two broad sets of data were generated numerically and analyzed. The first set consists of four fundamental total pressure distortions: once-per-rev circumferential, twice-per-rev circumferential, hub radial, and tip radial. The second set consists of three realistic turbofan distortion patterns: two analytic (though not fundamental) profiles, and one generated from S-duct computational results. For all investigated patterns, Radial Distortion Intensity and Circumferential Distortion Intensity are calculated in the manner described by ARP 1420, a guideline issued by the S-16 Turbine Engine Inlet Flow Distortion Committee for measuring total pressure distortions in turbomachinery. Additionally, interpolated total pressure contours are generated for each measurement configuration. These were then used to make point-to-point comparisons between the actual and estimated data. While total pressure distortion was used as the variable of interest for the majority of this paper, the conclusions may be applied to swirl, temperature, or any other flow property measured using a probe rake or traversing probe.

Abstract Achieving the objectives of the American and European long-term air transport in terms of emission reduction requires to shift to hybrid/electric aircraft by 2050. This implies the development of distributed propulsion systems using multiple low-diameter lightly-loaded shrouded rotors, operating at high rotational speed and moderate flight Mach number. The relevance of characteristic maps of such a propulsive system is discussed in this paper. It is based on an integrative approach, that takes advantage of the two different formalisms generally found when dealing with aeronautical propulsion : either propeller diagrams (external aerodynamics) or turbomachinery maps (internal aero-dynamics). It outlines the main parameters representing the shrouded rotors performance and the parameters that drive them in a self-similar way. This is done by dimensional analysis, taking into account the geometric and operating parameters of both the rotor and the shroud and their combined variables. The self-similarity is observed with potential flow and RANS computations, which credits the proposed formalism as a powerful tool for scaling the shrouded rotors.

Abstract The main intent of this work is the exploration of the rotor-only fan design-space to identify correlations between fan performance and enriched geometric and kinematic parameters. In particular, the aim is to derive a multidimensional “Balje chart”, where the main geometric and operational parameters are taken into account in addition to the specific speed and diameter, to guide a fan designer towards the correct choice of parameters such as hub solidity, blade number, hub-to-tip ratio. This multidimensional chart was built using performance data derived from a quasi-3D in-house software for axisymmetric blade analysis and then explored by means of machine learning techniques suitable for big data analysis. Principal Component Analysis (PCA) and Projection to Latent Structure (PLS) allowed finding optimal values of the main geometric parameters required by each specific speed/specific diameter pair.

Abstract The aim of the present paper is to improve the physical understanding of flow irregularities in the blade passing signal of turbomachinery rotors, since the novel stall warning method presented in part I is based upon those irregularities. For this purpose, a complementary instrumentation was used in a single stage axial compressor. A set of pressure transducers evenly distributed along the circumference surface mounted in the casing near the rotor tip leading edges is measuring the time-resolved wall pressures simultaneously to an array of transducers recording the chord-wise static pressures. The latter allows for plotting quasi-instantaneous 2D-pressure contours. Any occurring flow disturbances causing the before mentioned irregularity can later be classified using validated frequency analysis methods being applied to the data from the circumferential sensors. While leaving the flow coefficient constant, a continuously changing number of prestall flow disturbances appears to be causing the very spectral signature which is known from investigations on Rotating Instability. Any arising number of disturbances is matching a specific mode order to be found within the spectral signature. While the flow coefficient is reduced the propagation speed of prestall disturbances increases linearly as the speed seems to be independent from the clearance size. Data taken beyond the stalling limit demonstrate a complex superposition of stall cells and flow disturbances which the title “prestall disturbance” therefore doesn’t fit to precisely any more. Different convection speeds allow the phenomena to be clearly distinguished from each other.

Abstract A simple tip leakage flow (TLF) model which consists of a square duct with a longitudinal slit on the top of a side wall is proposed to reproduce the jet flow/main flow shear mechanism of the tip leakage vortex (TLV) rolling-up in turbomachinery. Large-eddy simulation (LES) is employed to investigate the turbulence characteristics of the flow model under low Reynolds number condition. The geometry and boundary conditions of the flow model are simplified from a compressor rotor and modified to apply to low-Re condition for LES. The vortex structures and turbulence characteristics of the LES results are compared with the measurements of the rotor. It is found that the flow model could reproduce similar flow field and turbulence structures compared with the TLF in the real rotor, thus it can be used to investigate the turbulence in practical flows. Reynolds-Averaged Navier-Stokes (RANS) calculations are also carried out. The mean flow and turbulence behaviors of different cases are analyzed. The budgets of turbulent kinetic energy ( k ) are analyzed to investigate the turbulence transport nature in the TLF model, indicating that the non-equilibrium transport process of k is significant, especially the pressure and turbulent transport, which is not predicted by RANS.

Abstract The unsteady flow prediction for turbomachinery applications relies heavily on unsteady RANS (URANS). For flows that exhibit vortex shedding, such as the wall-jet/wake flows considered in this study, URANS is unable to predict the correct momentum mixing with sufficient accuracy. We suggest a novel framework to improve that prediction, whereby the deterministic scales associated with vortex shedding are resolved while the stochastic scales of pure turbulence are modelled. The framework first separates the stochastic from the deterministic length scales and then develops a bespoke turbulence closure for the stochastic scales using a data-driven machine-learning algorithm. The novelty of the method lies in the use of machine-learning to develop closures tailored to URANS calculations. For the walljet/wake flow, three different mass flow ratios (0.86, 1.07 and 1.26) have been considered and a high-fidelity dataset of the idealised geometry is utilised for the sake of model development. This study serves as an a priori analysis, where the closures obtained from the machine-learning algorithm are evaluated before their implementation in URANS. The analysis looks at the impact of using all length scales versus the stochastic scales for closure development, and the impact of the extent of the spatial domain for developing the closure. It is found that a two-layer approach, using bespoke trained models for the near wall and the jet/wake regions, produce the best results. Finally, the generalisability of the developed closures is also evaluated by applying a given closure developed using a particular mass flow ratio to the other cases.

Abstract Strong hydrodynamic instabilities are often observed in rotating cavities of turbomachinery in the aerospace industry for specific thermal and operating conditions. This highly unsteady motion of the flow, often referred to as ‘pressure bands’ phenomenon can be detrimental and jeopardize the structural integrity of aeronautical engines by damaging turbomachinery components or turbopumps in the context of space applications. In order to avoid this fluid structure interaction, a deeper understanding of the stability of the flow of rotating cavities becomes compulsory. A few studies enabled to point out the spatial distribution and origins of such multi-frequency flow. However control solutions have not yet been set up. In this context, an academic stator/rotor cavity is studied in this paper with a sensitivity analysis to base flow modification based on linear stability analysis. This latter points out the regions of the stator boundary layer of each mode where the flow should be modified. Possible passive control strategies are then proposed from the interpretation of the results. The more relevant solution being suction and injection actuators directly introduced on the stationary disk of the cavity.

Abstract The rapid regulation of gas turbine power plants is becoming increasingly important. In the field of stationary gas turbine power augmentation droplet-laden flows receive special attention as they cool the compressor’s inlet temperature. High fogging is an effective and economic method that secures power supply by reacting rapidly to fluctuating power generation. However, water droplets effect the aerodynamic performance of the compressor negatively. Therefore, the Laboratory of Turbomachinery investigated the interaction between droplets and blades at midspan by using diverse experimental methods to comprehend aerodynamic performance and water film induced losses. Additionally, the knowledge of the separation’s extent for three-dimensional flows with corner separation is of great importance as it leads to a blockage of the flow cross-section and therefore contributes to the losses. This paper enlarges this knowledge base by studying the interactions of the dispersed phase with the gas flow in three-dimensional flow structures as they are found in the near wall regions of every turbo compressor. By utilizing a three-dimensional Phase-Doppler Anemometer spatially resolved information on velocity mean vector, its higher moments as well as droplet diameters were extracted from the separated region for a compressor blade, in middle cross section design, with spray-injection. Streamlines that derive from the mean velocities visualize key features such as saddle and focal points. Then an integral approach is presented to evaluate the effect of the dispersed phase on the flow structure compared to dry flow and the impact back on different diameter classes. Finally, the velocity in streamwise direction shows the influence of the disperse phase on the aerodynamic blockage.

Abstract This two-part paper presents the unsteady flow interactions between an engine-representative high-pressure turbine (HPT) and low-pressure turbine (LPT) stage, connected by a turbine center frame (TCF) duct with non-turning struts. The setup was tested at the high-speed two-spool test turbine facility at the Institute for Thermal Turbomachinery and Machine Dynamics at Graz University of Technology and includes relevant purge and turbine rotor tip leakage flows. Due to the complexity of such a test, the unsteady component interactions in an HPT-TCF-LPT module have not received much attention in the past and require additional analysis to determine new approaches for further performance improvements on the system level. The flow downstream of an HPT is highly unsteady and dominated by statorrotor interactions, which affect the flow behavior through the downstream TCF and LPT. To capture the unsteady flow structures, time-resolved aerodynamic measurements were carried out with a fast-response aerodynamic pressure probe (FRAPP) at three different measurement planes. In this first part of the paper, the time-resolved and phase-averaged flow fields with respect to the HPT and LPT trigger are studied. Since the two rotors are uncorrelated, the applied method allows the identification of the flow structures induced by either of them. Upstream of the LPT stage, the HPT flow structures evolving through the TCF duct dominate the flow fields. Downstream of the LPT stage, the flow is affected by both the HPT and the LPT secondary flow structures. The interactions between the various stator rows and the two rotors are detected by means of time-space plots and modal decomposition. To describe the fluctuations induced by both rotors, particularly the rotor-rotor interaction, the Rotor Synchronic Averaging (RSA) is used to analyze the flow field downstream of the LPT. The second part of the paper decomposes the flow fields to gain additional insight into the rotor-rotor interactions using the Proper Orthogonal Decomposition (POD) and RSA methods. The paper highlights the need to account for the HPT-induced unsteady mechanisms in addition to the LPT flow structures and the interaction of both to arrive at improved LPT designs.

Abstract Loss analysis is a valuable technique for improving the thermodynamic performance of turbomachines. Analysing loss in terms of the ‘mechanical work potential’ (Miller, R.J., ASME Turbo Expo 2013 , GT2013-95488) provides an instantaneous and local account of the thermal and aerodynamic mechanisms contributing to the loss of thermodynamic performance. This study develops the practical application of mechanical work potential loss analysis, providing the mathematical formulations necessary to perform loss analysis using practical Reynolds-Averaged Navier-Stokes (RANS) or Large Eddy Simulations (LES). The analysis approach is demonstrated using RANS and LES of a linear compressor cascade, both with and without incoming wakes. Spatial segmentation is used to attribute loss contributions to specific regions of the flow, and phase-averaging is performed in order to associate the variation of different loss contributions with the periodic passage of wakes through the cascade. For this un-cooled linear cascade, viscous dissipation is the dominant source of loss. The analysis shows that the contribution of the viscous reheat effect depends on the operating pressure of the compressor stage relative to the ambient ‘dead state’ pressure — implying that the optimal blade profile for a low-pressure compressor stage may be different from the optimal profile for a high-pressure compressor stage in the same engine, even if the operating conditions for both stages are dynamically-similar.

Abstract Performance and flow measurements are carried out to investigate the impact of varying the geometry of axial casing grooves on the stall margin and efficiency of an axial turbomachine. Prior studies have shown that skewed semi-circular grooves installed near the blade leading edge (LE) have multiple effects on the flow structure, including ingestion of the tip leakage vortex (TLV), suppression of backflow vortices, and periodic variations of flow angle. To determine which of these phenomena is a key contributor, the present study examines the impact of several grooves, all with the same inlet geometry, but with outlets aimed at different directions. The “U” grooves that have circumferential exits aimed against the direction of blade rotation achieve the highest stall margin improvement of well above 60% but cause a 2.0% efficiency loss near the best efficiency point (BEP). The “S” grooves, which have exits aimed with the blade rotation, achieve a relatively moderate stall margin improvement of 36%, but they do not reduce the BEP efficiency. Other grooves, which are aligned with and against the flow direction at the exit from upstream inlet guide vanes, achieve lower improvements. These trends suggest that causing high periodic variations in flow angle around the blade leading edge is particularly effective in extending the stall margin, but also reduces the peak efficiency. In contrast, maintaining low flow angles near the LE achieves more moderate improvement in stall margin, without the maximum efficiency loss. Hence, of the geometries tested, the S grooves appear to have the best overall impact on the machine performance. Velocity measurements and flow visualizations are performed in an axial plane located downstream of the grooves, near the trailing edge of the rotor. Reduced efficiency or performance co-occurs with elevated circumferential velocity in the tip region, but differences in the axial blockage are subtle. Yet, near the BEP, the regions with reduced axial velocity, or even negative velocity between the TLV and the endwall, are wider behind the U grooves compared to the S grooves. The vorticity profiles also show that at low flow rates the TLV is ingested entirely by the grooves, in contrast to the best efficiency point, where a considerable fraction of the TLV rollup occurs downstream of the grooves.

Abstract Modeling of the engine behavior in windmilling conditions is an important engineering objective. The relight capability of the engine is mainly associated with the air mass flow rate that passes through the engine high pressure core in those conditions. This is one of the parameters that drive the combustion chamber volume. Predicting the engine behavior is challenging, especially early in the development process. The pressure losses along the core are distributed between the different stages of the compressors and turbines, which are operated extremely far from their design point. Engine manufacturers must anticipate with sufficient margins to ensure that the specifications are met when the engine is finally qualified in flight. This article focuses on the behavior of compressor cascades in such conditions, corresponding to high negative incidences. A recently designed high pressure compressor is studied in windmilling conditions using three complementary approaches. First, engine tests are used to obtain validated 0D data for two flight conditions. Then, state of the art Lattice-Boltzmann Method (LBM) simulations are carried out to have insight in the flow characteristics inside the compressor. They are compared to the available experimental data. Finally, a 1D model using stage by stage Euler equation for turbomachinery is used. This kind of modeling is of particular interest because it can be used early in the design process. The correlations for losses and deviation angle from the literature are modified to account for the particularity of the flow in those conditions. One shows that the three approaches give consistent results.

Abstract A preliminary 2D numerical investigation of the active control of unsteady cavitation by means of one single synthetic jet actuator (SJA) is presented. The SJA has been applied to hinder the intrinsic instabilities of a cloud cavitating flow of water around a NACA 0015 hydrofoil with an angle of attack of 8° and ambient conditions. It has been placed inside the inception region at a distance of 16% of the chord from the leading edge. Concerning the numerical approach, a Eulerian homogeneous mixture/mass transfer model has been used, in combination with an extended Schnerr-Sauer cavitation model and a Volume of Fluid (VOF) interface tracking method. The synthetic jet has been modeled by means of a user-defined velocity boundary conditions based on a sinusoidal waveform. A sensitivity analysis has been first performed in order to evaluate the influence of the main control parameters, namely the momentum coefficient C μ , the dimensionless frequency F + and the jet angle α jet . By combining the cavitating vapor content and the impact on the hydrodynamic performance, the best performing SJA configuration has been retrieved. Then, a deeper analysis of the vapor cavity dynamics and the vorticity field has been conducted in order to understand the modification of the main flow produced by the synthetic jet. The best SJA configuration was observed at C μ = 0.0002, F + = 0.309 and α jet = 90°, which led to a reduction of both the average vapor content and the average torsional load in the measure of 34.6% and 17.8% respectively. A reduction of the average pulsation frequency of the pressure upstream confirmed the beneficial effect of the SJA. The analysis of the coupled dynamics between vapor cavity-vorticity and their POD-based modal structures highlighted that the benefit of the SJA lies on preventing the growth of a thick sheet cavity which tends to cause the development of the highly cavitating cloud dynamics after the cavity breakup. This is mainly due to an additional vorticity close to the hydrofoil surface just downstream the SJA, as well as a local pressure modification close the SJA during the blowing stroke.

Abstract Counter-rotating turbomachines have the potential to be high efficiency, high power density devices. Comparisons between conventional and counter-rotating turbomachines in the literature make multiple and often contradicting conclusions about their relative performance. By adopting appropriate non-dimensional parameters, based on relative blade speed, the design space of conventional machines can be extended to include those with counter-rotation. This allows engineers familiar with conventional turbomachinery to transfer their experience to counter-rotating machines. By matching appropriate non-dimensional parameters the loss mechanisms directly affected by counter-rotation can be determined. A series of computational studies are performed to investigate the relative performance of conventional and counter-rotating turbines with the same non-dimensional design parameters. Each study targets a specific loss source, highlighting which phenomena are directly due to counter-rotation and which are solely due to blade design. The studies range from two-dimensional blade sections to three-dimensional finite radius stages. It is shown that, at hub-to-tip ratios approaching unity, with matched non-dimensional design parameters, the stage efficiency and work output are identical for both types of machine. However, a counter-rotating turbine in the study is shown to have an efficiency advantage over a conventional machine of up to 0.35 percentage points for a hub-to-tip ratio of 0.65. This is due to differences in absolute velocity producing different spanwise blade designs.

Abstract Long-term efforts have been made to understand loss generation and its reduction in the field of axial turbomachines. The traditional approach to losses for an isolated blade row considers the profile and the secondary losses as a result of viscous flow. The additional kinds of losses in the stage are connected with the shear stress in the mixing process. These losses result from the mixing of the main stream flow with 1) the stator leakage injected through the root axial gap and 2) the return of the tip leakage over the bucket shroud. This article focuses on the first type of mixing losses. The leakage to the main stream flow ratio and the root reaction are the two key parameters investigated in this study. The primary data source for this study is the experiment. An experimental single stage air turbine was modified to set and precisely measure the stator leakage flow. Three configurations of the single-stage test rig with different reaction levels were tested. The second data source for this study is CFD computation. These computations are applied to different geometries and conditions from the experiment; they are derived from real steam turbine stages designed in DSPW. The computations simulate multistage configuration and real steam is considered as the working fluid. CFD computations were performed in the commercial software ANSYS CFX. Each configuration task was computed in three iterative steps. Each step takes the distribution of the flow parameters on the boundary domains from the previous iteration. The final results from this ‘repeating boundary conditions’ approach better correspond with the real expansion in a multistage configuration. The two data sources are not directly comparable. The experiment is used for validation of the trends. The computations provide the possibility of a multi-parametric study. The multi-parametric study is necessary to obtain a more general loss model which can be used during turbine design. The evaluation of the experimental and numerical parts focuses on a comparison of the overall stage performance. Stage efficiency and reaction are presented in relation to the ratio between leakage and main stream flow.

Abstract Modern turbomachinery faces increased performance demands in terms of efficiency, compactness, and pressure-rise. Advancements in computational technology have allowed numerical methods to become the backbone of design development efforts. However, the unique complexities of centrifugal compressor flow-fields pose difficult computational problems. As such, advanced experimental methods must be used to obtain high-quality datasets to further inform, improve, and validate computational methods in complex flow regimes. Recent experimental work on a high-speed centrifugal compressor has provided detailed, unsteady, three-component velocity data using Laser Doppler Velocimetry. A passage vortex is present and its nascent tied to the increased incidence at mid-span associated with impeller wake flow. This vortex begins in the hub-pressure side corner and grows to fill the passage and become temporally stable. The vortex development is unsteady in nature and the unsteady effects persist 40% downstream of the throat. Distinct jet and wake flow patterns from the impeller also do not agglomerate until 40% downstream of the throat. Additionally, the critical impact of the unsteady flow development on the time-averaged flow-field is explained.

Abstract The efficiency of modern axial turbomachinery is strongly driven by the secondary flows within the vane or blade passages. The secondary flows are characterized by a complex pattern of vortical structures that origin, interact and dissipate along the turbine gas path. The endwall flows are responsible for the generation of a significant part of the overall turbine loss because of the dissipation of secondary kinetic energy and mixing-out of non-uniform momentum flows. The understanding and analysis of secondary flows requires a reliable vortex identification technique to predict and analyse the impact of specific turbine designs on the turbine performance. However, literature shows a remarkable lack of general methods to detect vortices and to determine the location of their cores and to quantify their strength. This paper presents a novel technique for the identification of vortical structures in a general 3D flow field. The method operates on the local flow field and it is based on a triple decomposition of motion proposed by Kolář. In contrast to a decomposition of velocity gradient into the strain and vorticity tensors, this method considers a third, pure shear component. The subtraction of the pure shear tensor from the velocity gradient remedies the inherent flaw of vorticity-based techniques which cannot distinguish between rigid rotation and shear. The triple decomposition of motion serves to obtain a 3D field of residual vorticity whose magnitude is used to define vortex regions. The present method allows to locate automatically the core of each vortex, quantify its strength and determine the vortex bounding surface. The output may be used to visualize the turbine vortical structures for the purpose of interpreting the complex three-dimensional viscous flow field, as well as to highlight any case-to-case variations by quantifying the vortex strength and location. The vortex identification method is applied to a high-pressure turbine with three optimized blade tip geometries. The 3D flow-field is obtained by CFD computations performed with Numeca FINE/Open. The computational model uses steady-state RANS equations closed by the Spalart-Allmaras turbulence model. Although developed for turbomachinery applications, the vortex identification method proposed in this work is of general applicability to any three-dimensional flow-field.

Abstract One of the main goals for modern aircrafts is to lower the fuel consumption and noise emissions without worsening the aerodynamic performance. One possibility to lower the fuel consumption is to reduce the skin-friction losses of vanes and blades inside the engine. Therefore, this paper is about the aeroacoustical as well as the aerodynamical effects of a riblet foil applied on the suction side surface of turbine exit guide vanes (TEGVs) of a 1½ stage low pressure turbine (LPT). There have been numerous studies concerning riblets but none using them in a LPT. In general, if riblets are applied on the suction side of vanes or blades, they lower the drag and increase the lift. Test runs were performed under two different operating points in a subsonic test turbine facility for aerodynamic, aeroacoustic, and aeroelastic investigations (STTF-AAAI) located at the Institute for Thermal Turbomachinery and Machine Dynamics at Graz University of Technology. One operating point was the design point of the riblets and the second one an off-design point. During the test campaign, two different set-ups have been investigated. One configuration with riblets applied on the suction side of the TEGVs, and one configuration with a smooth foil on the vanes to achieve the same thickness as the first set-up. This smooth configuration serves as a reference case. The tested riblet structure was of trapezoid type with 45 μm tip distance and a height to tip distance ratio of 0.45. The acoustical data has been obtained by using flush mounted condenser microphones, rotated over 360 deg around the flow channel. The aerodynamical data was obtained by using an aerodynamic five-hole-probe as well as a trailing edge probe. Measuring in planes up- and downstream of each TEGV allowed the comparison of a rough pressure loss estimation between the two studied set-ups. The present work gives a closer insight into the change of the acoustical and aerodynamical behaviour by applying riblets to LPT vanes.

Abstract Pulsed detonation combustion is not a new research topic. However, since the detonation process was first observed in 1881, the interest in it grew substantially in the last decades. Because the gas turbines have reached their technological maturity, the scientific community has started looking into novel thermodynamic cycles, such as detonation-based cycles. Numerous studies have been recently published in the field of pulsed detonation combustion, both numerical and experimental and major breakthroughs have been achieved in understanding and controlling the phenomena. However, the topic remains of mostly academic interest, one of the reasons is that practical implementation of it is reliant turbomachinery that would efficiently convert the pulsed, high peak, pressure into useful work. The few studies conducted on classical, existing turbines, show an efficiency of around 50% when coupled with a PDC. The low efficiency has been directly connected with the shock wave losses. For this reason, the design of turbines with supersonic inlet, and associated performance assessments have been researched. This work, however, has supersonic steady inlet Mach number or sinusoidal pulsating conditions around an average subsonic value. No public literature exists on the performances of turbines operating at pulsating inlet conditions similar to the outlet of a PDC. The current paper tackles exactly this issue. The geometry for a turbine stator row was designed based on supersonic inlet design criteria. This geometry was then subjected to CFD numerical simulations. First, the pressure losses associated with a constant supersonic inlet were numerically determined to be a little over 26%. The next step was to assess the pressure losses of the same turbine row geometry in a transient approach. This time, the inlet conditions were set to be variable in time. The values were taken from a 1D in-house code computing the parameters at the outlet of a PDC working on hydrogen and air under stoichiometric conditions. This inlet conditions give a much better insight with respect to the flow within a turbine row when coupled with a PDC. It was observed that the pressure losses, computed as a time average for a period corresponding to the PDC functioning frequency were of 12%. This value is much less than that for a constant supersonic inlet, mostly due to the turbine being exposed to the shock waves for less time.