This paper describes the full-scale turbine section validation of the GT36 heavy duty gas turbine, which was conducted in a test Power Plant in Birr, Switzerland. The GT36 Test Power Plant is extensively instrumented with both standard and specialised instrumentation. In the turbine section, specialised instrumentation includes metal and gas thermocouples, thermal paint, pressure sensors, hot gas rakes, strain gauges, five-hole probes, pyrometers and tip timing sensors. Similar specialised instrumentation also exists for the compressor, combustor and the rotor sections. Three major test campaigns were conducted over an extended period, which consisted of both long and short duration tests, including a range of off-design tests. Within the turbine section, detailed transient and steady-state measurements were obtained of the stage inlet pressures and temperatures, airfoil surface pressures and metal temperatures. These measurements indicated that both the aerodynamic and cooling performances of the turbine blades and vanes are highly consistent and repeatable over a range of operating conditions. Detailed comparisons of the measured engine pressures and temperatures with predictions also indicated that there was generally a very good match in the Mach numbers and metal temperatures for all the turbine blades and vanes.

In the process of turbine acceleration, due to the influence of compressor and complex secondary air system, the change process of coolant purge flow is relatively lagging behind that of mainstream flow and rotational speed. The lagging egress of coolant flow influence the aerodynamic performance and endwall cooling effectiveness of turbine acceleration process. The flow field and aerothermal performance of two-stage axial turbines combined with rim seal structures and coolant purge flow lagging effects in the turbine acceleration process was numerically investigated using Unsteady Reynolds-Averaged Navier-Stokes (URANS) via SST turbulence model. The effects of lagging coolant purge flow across the rim seal on the turbine aerodynamics and endwall cooling effectiveness were analyzed. The obtained results show that the turbine aerodynamic efficiency obtains the maximum value when the coolant purge flow lagging time equals to half the acceleration time at the same rotational speed after the end of lagging times. The total-to-total efficiency for the second stage is more sensitive to lagging times. The turbine output power is almost un-changed due to combination of additional work capacity and aerodynamic loss with the introduction of coolant. The turbine endwalls have the maximum averaged cooling effectiveness in the turbine acceleration process without consideration of the coolant purge flow lagging time. And endwall cooling effectiveness decreases with the increase of coolant purge flow lagging time at the same rotational speed and mainstream flow conditions. The detailed flow field of two-stage turbine considering interaction between the coolant purge flow and mainstream was also discussed. The present work provides the reference for the match design between the turbine mainstream flow and secondary air flow system.

In case of stationary power production, the matching between the turbocharger and the engine can be optimized around a specific operating point, rather than throughout a wide range like in automotive applications. This, in turn, may enable a finer optimization of the aerodynamic efficiency in comparison to existing industrial standards. The present study then reports a joint research activity aimed at defining an integrated approach to optimize the overall performance of an engine for stationary power production via a dedicated optimization of the turbocharger unit. In further detail, the aerodynamic optimization of the compressor, the turbine and the static parts is carried out by means of high-fidelity CFD simulations, while a 1D model of the powertrain is realized by means of the GT-Power software. An iterative procedure between the two simulation frameworks is established to account for the fluid dynamic connection between the engine and the turbocharger. Even though the selected case study was a “high quality” commercial product, the study shows how notable aerodynamic improvements can be achieved on turbocharger aerodynamics, thanks to the fact that the rangeability can be narrowed. Efficiency increases of 2.4% and of 3.4% have been reached for the compressor and the turbine, respectively. In the selected case study, this reflected into a moderate overall engine improvement (+0.5%), which however is not negligible in case of engine for stationary power production. Anyhow, the results suggest more promising prospects in case the methodology is applied since the early design of the engine.

Sweep in a transonic fan is conventionally used to reduce design point losses by inclining the passage shock relative to the incoming flow. However, future low pressure ratio fans operate to lower Mach numbers meaning the role of sweep at cruise is diminished. Instead, sweep might be repurposed to improve the performance of critical high Mach number off-design conditions such as high angle of attack (AOA). In this paper, we use unsteady computational fluid dynamics to compare two transonic low pressure ratio fans, one radially stacked and one highly swept, coupled to a short intake design, at the high AOA flight condition. The AOA considered is 35°, which is sufficient to separate the intake bottom lip. The midspan of the swept fan was shifted upstream to add positive sweep to the outer span. Based on previous design experience, it was hypothesised the swept fan would reduce transonic losses when operating at high AOA. However, it was found the swept fan increased the rotor loss by 24% relative to the radial fan. Loss was increased through two key mechanisms. i) Rotor choking: flow is redistributed around the intake separation and enters the rotor midspan with high Mach numbers. Sweeping the fan upstream reduced the effective intake length, which increased the inlet relative Mach number and amplified choking losses. ii): Rotor-separation interaction (RSI): the rotor tip experiences low mass flow inside the separation, which increases the pressure rise across the casing to a point where the boundary layer separates. The swept fan diffused the casing streamtube, causing the casing separation to increase in size and persist in the passage for longer. High RSI loss indicated the swept fan was operating closer to the rotating stall point.

Aerodynamics phenomena in compressors are highly unsteady and turbulent. Selecting a proper turbulence-modeling method is significant to reveal the complex flow mechanism in turbomachines. In the current paper, the shear stress transport (SST) model as an unsteady Reynolds-averaged Navier-Stokes (RANS) method, the scale-adaptive simulation (SAS) model, and the zonal wall-modeled large eddy simulation (zonal-LES) as two hybrid scale-resolving simulation approaches have been compared. These turbulence-modeling methods were employed to simulate a single rotor of a low-speed research compressor featuring a tip clearance of 1.3% of chord length. Comparisons were made between the simulation results and the experimental data at three operating points, and the flow fields at the design point have been specifically discussed in detail. The results show that the advantage of the zonal-LES model becomes obvious as the compressor throttles. The zonal-LES model brings a significant improvement over both the SST model and the SAS model in capturing the experimental data, especially the velocity distribution in the low-span region, as well as the loss near the endwalls. The SAS model as a scale-resolving method presents no benefits in predicting the relevant flow compared with the SST model, as the activation of the SAS source term is limited for this test case. For the loss prediction, the variation in the upper half-span region is mainly due to the different results in modeling turbulent characteristics of the tip leakage flow, whereas the mechanism behind the higher loss at the lower half-span predicted by the zonal-LES model is a consequence of the complex topology of the corner separation and the intensive mixing.

Civil air traffic is predicted to further grow in the near future. Hence, the development of aeronautical combustors will face major challenges to meet future stringent environmental regulations. In the present study, an innovative gas turbine combustor with angular air supply called Short Helical Combustor (SHC) is investigated. The main feature of this concept is the helical arrangement of the fuel injectors around the turbine shaft. Aiming at the implementation of a lean-burn concept, a low-swirl lifted flame is adopted. This flame is lifted off and not anchored to the injector which opens the potential of low NO x emissions due to a high degree of premixing within the combustor. In this work, isothermal flow characteristics of such a generic SHC combustor are studied by use of RANS predictions with special emphasis on the interaction of adjacent low-swirl flows. For evaluating the influence of injector parameters on the flow field, a parametric study based on single sector simulations is performed. It is shown that the asymmetrically confined swirling jet flow is strongly deflected towards the sidewall of the staggered SHC dome. The deflection of the flow is associated with an asymmetric pressure field in the vicinity of the burner which is generally known as Coandă effect. As a consequence of the deflected flow, the angular momentum flux at combustor outlet is increased. The interaction of the low-swirl jet and the SHC sidewall is investigated with regards to backflow momentum and residence time in the recirculation zone. It is concluded that by modifying the momentum of the air flow through the injector, the amount of recirculating air flowing back along combustor walls is strongly affected. The present work establishes an understanding of the underlying aerodynamics of the SHC concept which is essential for matching the requirements of lean lifted flames.

With a special focus on the industrial feasibility and the manufacturability, a recently proposed novel approach to centrifugal impeller blade inclination is adopted and investigated through extensive CFD analysis. The fan blades, originally aligned perpendicular to the impeller backplate, are inclined in either forward or backward direction. For the presented study, an industrially proven fan design is chosen for testing. Compared to the original design, the inclined fan blades possess an increased total blade area and at the same time providing variable inflow angles at the leading edges of the blades. These two factors are expected to alter the fan characteristic curves in providing an increased range of optimum performance while maintaining high aerodynamic efficiency. The results obtained show a clear trend in aerodynamic performance with the degree of inclination, where the characteristic curves rotate at about the design point, allowing local improvements either at overload conditions or part-load conditions of the fan. Moreover, the trends obtained show the tendency to agree well with the rudimentary models published in previous studies, even though it appears to be affected by the fan volute and the point of operation as well.

State-of-the-art axial compressors of gas turbines employed in power generation plants and aero engines should have both high efficiency and small footprint. Thus, in many cases, axial compressors are designed to have thin rotor blades and stator vanes with short axial distances, and are driven at high rotational speed. Recently, problems of high cycle fatigue (HCF) associated with forced response excitation have gradually increased as a result of these trends. Rotor blade fatigue can be caused not only by the wake and potential effect of the adjacent stator vane, but also by the stator vanes of two, three or four compressor stages away. Thus, accurate prediction and suppression method of them under the resonance condition are necessary in the design process. Concerning the forced response excitation associated with the adjacent stator vanes, there are many previous studies on simulating the vibration by fluid structure interaction (FSI) simulation. In these studies, the aerodynamic force acting on the blade is simulated by an efficient unsteady computational fluid dynamics (CFD) method such as the nonlinear harmonic (NLH) method. These methods can be available in commercial CFD solvers and can significantly reduce computational cost. However, there are few examples of the problems associated with the stator vanes from two and three compressor stages away and no efficient simulation method is available. In this study, the problem of rotor blade vibration caused by the stator vanes of two and three compressor stages away is studied. Ways to accurately predict and effectively control the vibration are also investigated. In the first part of the study, one-way FSI simulation is carried out using a full annulus CFD model. To validate the accuracy of the simulation, experiments are also conducted using a gas turbine test facility. The vibration level of the blade is measured using a blade tip timing (BTT) measurement system and the obtained results are compared with the simulated data. It is found that one-way FSI simulation can accurately predict the order of the vibration level. In the second part of the study, a method of controlling the forced response excitation is investigated by optimizing the clocking of the stator vanes. It is confirmed that by controlling the clocking of the stator vanes, the vibration amplitude can be effectively suppressed without reducing the compressor performance. Through this study, ways to evaluate and control the unsteady pressure force and vibration response of the rotor blade are validated. By optimizing the clocking of stator vanes, the blade vibration level can be effectively reduced.

Exhaust gas of an automotive fuel cell is enriched with water vapour and has a pressure potential which can be utilized by a turbine. The gas expansion in the turbine leads to droplet nucleation and condensation. This results in a release of latent heat and a decrease of the gaseous mass flow which has a considerable influence on the turbine performance. This study aims to numerically investigate the influence of these phenomena on the performance map of the radial turbine of an automotive fuel cell turbocharger. For this purpose, the classical nucleation theory and Young’s droplet growth law are integrated into an Euler-Lagrange approach. The results show an almost linear relation between the pressure ratio and the condensation while the specific aerodynamics of an operating point has only a minor influence. At 80 % relative humidity of the inflow, the investigated turbine showed condensation above a total-to-static pressure ratio of 1.8. Condensation leads to thermal throttling of the turbine and to a temperature increase of the rotor outflow of up to 50 K. Increasing humidity of the inflow increases the power output, but condensation losses reduce the efficiency.

Future turbofan engines seek shorter intakes to reduce the cruise fuel burn of a low pressure ratio, large diameter fan. However, shorter intakes increase the level of flow distortion entering the rotor when the aircraft angle of attack (AOA) is high, reducing thrust when critically needed. This paper considers how the fan rotor radial pressure ratio distribution and tip velocity triangle can be designed to improve thrust at high AOA. Full annulus, unsteady CFD is performed on three rotor designs coupled to a short intake. We show that rotor design for high AOA should be guided by three flow mechanisms. Mechanism i) is caused by high Mach number flow over the bottom intake lip, which chokes the rotor leading to high loss. Mechanism ii) is the loss generation in the rotor tip as it passes through an intake separation. Mechanism iii) shows radial flows through the rotor change both the amount and the way work is imparted on the flow. Two comparable rotor design philosophies for high thrust are proposed; high work or low loss. Rotors designed to a mid-high radial pressure ratio distribution impart high work on streamlines that migrate radially towards the hub and exit the rotor at highly cambered sections. Meanwhile, tip-high designs reduce choking losses in the midspan when operating with a separated intake, particularly when the tip velocity triangle is designed to high axial velocity diffusion over high camber. However, such designs suffer with higher tip losses after exiting an intake separation.

The S-shaped duct which transfers flow from the low-pressure fan to the engine core of modern large civil turbofans presents a challenging design problem. Aerodynamically it must accommodate a spatially and temporarily non-uniform inlet in conjunction with a complex flow field which will develop under the combined influence of pressure gradients and streamline curvature. It must also allow for the transfer of structural loads and services across the main gas path. This necessitates the use of structural vanes which can compromise aerodynamics, introduce unwanted component interactions and erode performance. Furthermore, this must all be achieved with minimum length/weight and without flow separation. This paper presents a comprehensive aerodynamic evaluation of three options for a low-pressure compressor transition duct containing (i) a long-chord, structural compressor outlet guide vane, (ii) a more aerodynamically optimal but non-structural outlet guide vane in conjunction with a small number of discrete radial load bearing struts and (iii) a fully integrated outlet guide vane and strut design. Evaluation was performed using a fully annular, low-speed test facility incorporating a 1 ½ stage axial compressor and transition duct representative of an engine design. Aerodynamic data were produced from miniature five-hole probe area traverses conducted at several locations with compressor/duct. The data suggest that all the options were viable. However, the aerodynamic vane and discrete struts produced the lowest system loss with the other two options being comparable. The performance of the structural vane was seen to be sensitive to off-design conditions producing a notably increased loss at a low flow coefficient. The more optimized aerodynamic vanes were much less sensitive to off-design conditions whilst the fully integrated design showed only very small changes in loss.

Centrifugal compressors employed in the oil and gas industry are operated at high gas pressure conditions and are used in a wide operation range. Accurate prediction of the rotating stall and the destabilizing aerodynamic force is one of the key technologies for these compressors, because rotating stall can sometimes cause severe problems with subsynchronous shaft vibration and limit its operation range. Thus, the aim of this study is to establish a method of accurately predicting the inception of rotating stall and its effect on shaft vibration. To achieve this, numerical investigations are carried out by unsteady Reynolds-averaged Navier-Stokes (RANS) simulation with a full annulus model of the compressor stage. Also, to assess the accuracy of the simulation qualitatively and quantitatively, a high-pressure compressor test rig that contains a shrouded impeller and a vaneless diffuser is built. To investigate the effect of the rotating stall on the shaft vibration, an experiment is carried out at relatively high gas pressure with the inlet pressure level exceeding 30 barA. In the first part of the study, the accuracy of compressor performance prediction is studied by steady computational fluid dynamics (CFD) simulation. It is found that by taking the wall roughness effect into account, the predicted performance shows good agreement with the experimental result. Thus, a subsequent study of the rotating stall is also carried out by considering its effect. In the second part of the study, the accuracy of predicting the rotating stall is studied. In the experiment, two types of rotating stall are measured. One is a multiple-cell stall induced in the vaneless diffuser, for which the speed of rotation is relatively low and the other is a one-cell stall induced in the impeller region. The properties of the multiple-cell stall agree with the previous experimental and numerical studies, and the rotating stall has the limited effect on shaft vibration. Conversely, the one-cell stall shows severe subsynchronous vibration. In this study, both types of stall prediction are examined by CFD simulation. It is found that the simulation can predict the inception of the rotating stall with relatively high accuracy as the predicted results show good agreement with the experimental results in terms of cell count, rotation speed, pressure fluctuation level and the effect on shaft vibration. Through this study, the effectiveness of unsteady CFD simulation is confirmed for these types of stall and vibration prediction.

Radial outflow turbines are an alternative for axial turbines for example in heat recovery applications. They are, however, also often characterized by ultra-low aspect ratios. In these designs, the secondary losses dominate the overall loss share, and under a certain aspect ratio, the secondary structures from the hub and shroud begin to interact. This interaction causes a decrease in aerodynamic performance. Previous studies have suggested that the general flow phenomena between radial outflow and axial turbines could share several similarities due to observed trends in performance prediction. The blade outlet Mach number is known to affect the spanwise positions of the secondary vortices in axial turbine blading and therefore, its effect is also tested here for an ultra-low aspect ratio radial outflow turbine cascade. In addition, there are currently no cascade level experimental data publicly available, and the suitability of axial turbine loss correlations under these conditions remains an open question. From this background, the current study presents an experimental, numerical, and loss correlation analysis of the effects of an isentropic Mach number in a radial outflow turbine cascade. An experimental campaign is used to validate the numerical model both quantitatively and qualitatively. In addition, the validity of the axial turbine loss correlation is extended to ultra-low aspect ratios by introducing a new variable called penetration length. The main findings are: 1. The flow phenomena do not differ significantly from what has been observed with axial turbines, 2. The effect of penetration length calculation method on the loss breakdown is relatively low, and 3. With ultra-low aspect ratio radial outflow turbines, the loss breakdown is markedly changed when the extended Benner’s approach is employed.

Film cooling is commonly used to protect turbine vanes and blades from the hot gases produced in the combustion chamber. The design and optimization of these systems can however only be achieved if a precise prediction of the fluid mechanics and film efficiency is guaranteed at a level where induced losses are fully mastered. Such a prerequisite induces at the numerical level to be able to identify and assess losses. In this context, the present study addresses loss assessment in a wall-resolved Large Eddy Simulation (LES) of the film-cooled high-pressure turbine blade cascade T120D from the European project AITEB II. The objectives are twofolds: (1) to evaluate the capacity of LES to predict adiabatic film cooling effectiveness in a mastered academic case; and (2) to investigate loss generation mechanisms in a fully anisothermal configuration. When it comes to LES predictions of T120D, the flow structure around the blade and the coolant jet organization are coherent with literature findings. Satisfactory agreements are furthermore retrieved for the pressure load prediction as well as the adiabatic film effectiveness if compared to the experiment. Loss generation is then investigated illustrating the fact that aerodynamics losses dominate mixing losses which are mainly located in the coolant film. This is in line with the temperature difference between the hot and coolant flows that is low for this experimental condition. Distinct contributions can however be made available by studying the local loss generation maps by means of Second Law Analysis if recast in the specific context of anisothermal flows when simulated by LES.

The purpose of this paper is to introduce the basics of life cycle cost analysis for use in an undergraduate, senior-level capstone, gas turbine engine design course. This paper will support the heightened interest within the military acquisition community that now requires life cycle cost analysis to be included in the proposals submitted by defense contractors. The capstone design course includes both the gas turbine engine cycle selection and engine component design that supports a particular aircraft application. While the students have been taught how to estimate the fuel costs, engine development costs, and the time-varying production costs of engines, they have not yet been provided instruction on how to factor all three types of costs into an engineering economics, time-value-of-money, present value analysis. This paper will fill that gap and serve as a resource for the students who must now consider life cycle cost as an element in their design decision matrix along with engine performance, technical risk, and development time. The typical case compares an engine where the upfront development and production costs associated with a more advanced level of technology are high early on in the life cycle but over time has a lower fuel cost compared to an engine with a lower development and production cost but with a higher fuel cost. This paper illustrates how the aerodynamics, thermodynamics, and engineering economics can be brought together to inform and defend a decision about which of the two (or more) alternatives is best. The engineering economic analysis is spreadsheet based and uses inflation adjusted, total annual costs to calculate the present value for use in a decision matrix.

The most relevant quality key numbers for the largest and most efficient Gas Turbine Combined Cycles (GTCC) are not (only) the data published by the original engine manufacturers OEM’s. Additional numbers are here evaluated with educated guesses based on published data of the latest announcements of the “big four OEM’s” [8]. Such data are of interest for potential customers but also for nailing down the current state-of-the-art for all kind of further cycle studies using turbomachinery components and also as a contemporary history record. Making educated guesses means thermodynamic 1D simulation based on additional assumptions for pressure losses and other cycle data, which have a limited influence on the (unpublished) target quality numbers, such as: • Mixed turbine inlet temperature Tmix. This is a key value describing the technology level. It can be derived independently of the (unpublished) TCLA value. It is a quality number for the general cooling design and for the secondary air systems. • Polytropic efficiency of the compressor blading. This number describes the aerodynamic quality of the compressor blading. • Polytropic efficiency of the turbine blading. It describes the quality level of both the blading aerodynamics and of the open air cooling design. • Distribution of the exergy losses within the GT and in the bottoming cycle. The exergy losses describe the remaining opportunities for further improvements in the thermodynamic cycle design. But they also indicate its limits. However already the determination of the Tmix is tricky. It depends on the analysis method and on the fluid data applied. The polytropic efficiency of the turbine blading and the exergy losses will depend both on the used methods and on the Tmix found. Achieving a trustable result therefore requires a transparent and reproducible method. In case of application of the found results for performance prediction of similar cycles the same method has to be applied in order to avoid mistakes. In this paper real gas data with consideration of dissociation in equilibrium are used, while the polytropic efficiencies are determined with an incremental method based directly on the classic definitions of Stodola [3] and Dzung [4]. Therefore the still most used method using semi-perfect gas properties and corresponding formulas is bypassed. In order to keep it as simple as possible the evaluation is limited to base load at ISO ambient condition (15°C, 60% relative humidity, sea level). The fuel is limited to pure methane according to the practice in current catalogue data. The main focus is on the gas turbine with its components. The steam bottoming cycle is captured with its effect on the overall exergy and energy balance of the GTCC, which identifies exhaust and condensation losses.

The effect of hydrogen enrichment of a premixed hydrogen-methane-air jet in hot vitiated crossflow was studied at atmospheric condition. The hot turbulent vitiated crossflow is generated by a symmetric array of 4 × 4 jet flames burning a lean mixture of natural gas and air in fully premixed condition at equivalence ratio φ cf = 0.7 and total thermal power of 50 kW. This crossflow is then used to ignite the premixed perpendicular jet of hydrogen-methane-air at ambient temperature. Three jet parameters are varied to study the effect of hydrogen addition on the flame morphology and stabilization mechanism: the hydrogen mass fraction of the H 2 /CH 4 fuel blend (ξ = 0 – 100%), the jet equivalence ratio (φ = 0.8 – 2.0) and the jet-to-crossflow momentum ratio (J = 3 – 12). High-speed hydroxyl (OH) chemiluminescence is used to obtain the time-resolved imaging of the reactive jet and to compute its time averaged morphology. OH planar laser induced fluorescence (OH-PLIF) is used to acquire OH concentration fields at the jet center plane. The jet morphology is analyzed by considering its mean trajectory, extracted from the experimental data and fitted with empirical correlations available from the literature. New correlations are proposed for the flame length, width and center of gravity as function of the hydrogen content. It is shown that with increasing hydrogen fraction, the flame is shortened and more compact, and it stabilizes close to the jet root. Another finding of this work is the reattachment of the flame at the base of the windward jet shear layer when hydrogen fraction is increased. Robust flame anchoring is observed for H 2 mass fractions of the CH 4 /H 2 fuel blend that exceed 50%. Moreover, it is shown using instantaneous OH-PLIF images that for these conditions of increasing hydrogen concentration, the windward shear layer features larger-scale coherent structures that govern the aerodynamics of the reactive premixed jet in turbulent vitiated crossflow.

Environmental requirements have led the air transportation industry to work towards reducing greenhouse gas emissions and mechanical noise levels. Nowadays, this sector contributes with 2% of the total greenhouse gas emissions, and there is a demand from global aviation regulators for further reducing this percentage. In the last years, the development of Hybrid-Electric Propulsion Systems (HEPSs) has grown. The HEPS combines an Internal Combustion Engine (ICE), for example, Gas Turbine (GT) or reciprocating engine, with an Electric Motor (EM), combining the inherent advantages of both. HEPSs present increased efficiency and operating safety in comparison with conventional ICE-powered systems. Furthermore, they can supply the electrical devices with power. This area of study is multidisciplinary in nature and, therefore, poses research challenges on ICEs, EMs, electronic converters, propeller design, monitoring and control systems, management and supervision systems, energy efficiency and optimization, aerodynamics and aircraft mechanical design. A research project aimed at the characterisation of hybrid-electric aircraft propulsion systems, and the construction of a HEPS prototype, is underway in Brazil. The system is essentially composed of a GT, an EM, three electronic converters, a battery bank and a propeller. It can operate with three different topologies: series, full-electric and turbo-electric. A test bench with all the necessary peripheral and analysis infrastructure is under construction. Present work aims to: (i) develop simplified models for all the test bench components, (ii) given a mission profile, show the results of an initial energy management computing code that determines the optimal hybridization strategy, and (iii) simulate various operating alternatives for the chosen mission profile. The results (i) highlight the impact of critical characteristics of the batteries on the HEPS performance, and (ii) demonstrate the application of the management code on optimizing the aircraft energy consumption for a given mission profile.

Surface roughness is well known to significantly influence turbine aerodynamics and heat transfer; different studies have been undertaken in the last decades, in order to precisely characterize its effects and pursue a reliable and unified CFD modelling approach. Despite the effort, further research is still required to completely fulfill the goal, due to the complexity of the considered environment, with many other aspects and flow characteristics factoring into the final behavior. In this work an experimental campaign was carried out to evaluate the heat transfer coefficient on a linear nozzle guide vane geometry. The adopted geometry has been developed and tested, at different inlet turbulence intensity, Reynolds and Mach number, at Von Karman Institute. The results achieved on a test article with smooth surface were made available. In the present work the effect of increased turbulence level and surface roughness were taken into account, respectively using passive grids and conditioning the test sample surface finishing. Experiments were conducted using a transient technique by measuring the surface temperature evolution by IR thermography. The collected results integrate the existing database available in the open literature in order to support development and benchmarking of numerical approaches aimed at a reliable characterization of these aspects.

While modern engine manufacturers devote significant efforts to the development of reliable and efficient machines, the introduction of novel, optimized components in the hot gas path represents a risky opportunity. Accurate experimental and numerical data are critical to assess the impact of new technologies on the harsh engine environment. The present study addresses the impact of a selection of high-performance rotor blade tips on the aerodynamic and heat flux field of a high pressure turbine (HPT) stage. A combined numerical and experimental approach is employed to characterize the interaction of the tip leakage flow with the rotor secondary flows and the casing heat transfer mechanisms for each individual tip geometry. The turbine stage is tested at engine-scaled conditions in the rotating turbine facility of the von Karman Institute. For the present study, the turbine rotor is operated in rainbow configuration to allow the simultaneous testing of multiple blade tip geometries. RANS simulations are employed to predict the aerodynamic and thermal field of the individual profiles using test-calibrated boundary conditions. Isothermal steady computations are performed at different wall temperatures to compute the adiabatic wall temperature and heat transfer convective coefficient. Low-order models are used to represent the over-tip thermal field and the driving heat transfer mechanisms. The time-resolved outlet flow is characterized using a vortex tracking technique and high frequency aerodynamic measurements to identify the rotor secondary flow structures.