Over the service time, the rotating parts of turbine engine vary in their geometry. When aircraft take off or fly through a volcanic ash cloud, the particles are sucked into the engine, impinge the blade and gradually erode the surface. The impinging between particles and blades is responsible for the increase of the surface roughness. Also, during the long-time operation, the function of the blade’s stacking law combined with the centrifugal force could cause deviation of the stagger angle. Moreover, blade tip clearance could vary because of the casing deformation. All the deformation of geometry could severely reduce the engine performance and thus engine life. The work presented in this paper focused on the influence of geometry deformation in a real low-pressure compressor. The investigation is more difficult than most of the previously published researches with a total of five stages being considered. Due to the irregularities in geometry, it is difficult to numerically assess the performance of the compressor. The aim of this study is to give an analysis method that allows an efficient and accurate estimation of the performance for multistage compressor with geometry deformation. In the first step, the geometry models with different deviation in tip clearance, roughness and stagger angle were established respectively. A CFD study was then applied to the compressor with RANS method to calculate the flow field with different types of deformation. The variation of overall performance due to the deformation was finally analyzed to identify the dominant factor on influencing the performance of the compressor among different types of geometry deformation. A method based on polytropic efficiency analysis and flow field analysis was also established to specifically analyze which stage is most sensitive to the geometry deformation. The results show a significant influence of geometric deformation on the efficiency, total pressure rise and flow range of the multistage compressor. The conclusions of this study would provide an important guidance for engine overhaul in the factory.

The accurate prediction of wall–roughness effects in turbomachinery is becoming critical as turbine designers address airfoil surface quality and degradation concerns arising from the shift to advanced ceramic matrix composite (CMC) or additively–manufactured airfoils operating in higher temperature environments. In this paper, a recently developed computational capability for accurate and efficient scale–resolving simulations of turbomachinery is extended to analyze the boundary–layer separation and transition characteristics in a rough–wall low–pressure turbine (LPT) cascade. The computational capability is based on an entropy–stable discontinuous–Galerkin spectral–element approach that extends to arbitrarily high orders of spatial and temporal accuracy, and is implemented in an efficient manner on a modern high–performance computer architecture. Results from the scale–resolving simulations of both smooth and rough airfoil cascades are presented and compared to previous experiments and numerical simulations. The results show that the suction–surface boundary layer undergoes laminar separation, transition, and turbulent reattachment for the smooth airfoil cascade, while in the presence of roughness the separation and transition behavior of the suction–surface boundary layer is substantially modified. The differences between the smooth and rough airfoil cascades are then highlighted by a detailed analysis of their respective turbulent flow fields.

In the context of smooth surfaces where no industrial process modifies the flow and where no roughness affects the boundary layer flow, there are configurations today where the correct heat flux prediction is still unattained for certain operating points. This is the case of the LS89 configuration that has shown to be of great difficulty to accurately simulate the thermal fields for high Reynolds number flows even when performing wall-resolved Large Eddy Simulations (LES). The physics of the studied operating point (MUR235) are especially complex due to the interaction of a transitioning boundary layer, shock waves and free-stream turbulence injected at the inlet. In this paper, free-stream turbulent specifications are seen to be important towards the capture of the heat transfer profile on most regions of the blade. The boundary layer is found to be transitional when either artificially raising the level of turbulence at the inlet or by using a highly refined mesh that induces the generation of turbulent spots that increase the heat transfer. The important refinement done improves the heat flux predictions to the point it is approaching the experimental data.

Compressor fouling is a severe problem for both heavy-duty and aero-propulsion gas turbines. Particles can impinge on the blade and annulus surfaces, sticking there. The consequences of particle deposition are the increase of the roughness and an uncontrolled variation in the surface shape. These problems have a major effect on the performance of the compressor over time. Variations in the flow field can make the flow quantities close to the deposit to change, and it may happen that the conditions for the sticking do not hold any longer. If this is the case, the build-up detachment may happen. This occurrence can mitigate the fouling effects and may be exploited for keeping the performance of the compressor as high as possible over the operating period. In this work, an innovative model is proposed in order to evaluate the adhesion forces and the possible detachment. Particularly, the same forces that keep a gecko stuck to a surface are considered: the van der Waals forces (due to the proximity of the two bodies) and the Laplace force (due to the curvature of the liquid film related to the humidity). The so formulated model, named gecko-like for such a reason, is used for the numerical analyses of a deposition problem. Both the sticking and possible build-up detachment are considered. The outcome of this work can be regarded as an a-priori estimate of the forces to be kept into account when dealing with compressor fouling problems.

Solid particle ingestion is one of the main compressor degradation mechanisms for both heavy-duty and aero-propulsion gas turbines. Particles impacting the inner surfaces of the machine can stick there forming deposits. The presence of such deposits reflects on the reduction in performance of the machinery. Over last years, several methods have been developed in order to study the problem from the numerical standpoint. Examples of these techniques are the mesh morphing approach and the added-roughness-and-thickness method. In this work, an innovative procedure is proposed in order to evaluate the losses and the variation in the fluid flow due to the deposits. Particularly, an algorithm capable of determining the microscale deposition pattern has been developed. By using this methodology, a comprehensive analysis of the variation of the performance of the compressor over time can be carried out. The deposition severity and the subsequent roughness variation can be kept into account in a very detailed and precise fashion. Furthermore, this approach overcomes the difficulties that may arise by using a mesh morphing technique. The computational grid is not modified and thus its quality is retained, without re-meshing requirements, even for large deposits. The local roughness variation is accounted for without extra-effort. The procedure developed, shown here in deposition problems, can be easily extended to erosion or even icing problems. The only parameter to be changed is the model that takes care of the particle-wall interaction, using an erosion rather than an icing law.

Laser-Powder Bed Fusion (L-PBF) is an additive manufacturing technique used to melt metal material into solid three-dimensional parts. While offering a high degree of design freedom, L-PBF still has technical restrictions, like the achievable surface roughness, resolution and the need for support structures in overhanging areas. [1] Currently, L-PBF is used mainly to produce small batches of parts and prototypes. [2] In order to fully industrialize the technology, the research campus in Aachen is investigating possible future applications in turbomachinery while developing the corresponding processes with industry partners. Sealing systems, like honeycomb seal strips in gas turbines often require time-consuming joining and assembly operations that can be avoided by building up the structure monolithically using L-PBF. The following process development study proves the feasibility of manufacturing honeycombs with L-PBF using the Nickel-based super-alloy Inconel 718 (IN718) on an EOS M290 machine. Here, we have evaluated the economic aspects of different build orientations of the seal strips. Afterwards, we conducted a systematic parameter study with continuous and pulsed wave laser emission and investigated the resulting wall thicknesses. A reduction in wall thickness of about 30% can be observed when a modulated laser is used.

Fouling affects gas turbine operation and airborne or fuel contaminants, under certain conditions, become very likely to adhere to surfaces if impact takes place. Particle sticking implies the change in shape in terms of roughness of the impinged surface. The consequences of these deposits could be dramatic: these effects can shut an aircraft engine down or derate a land-based power unit. This occurrence may happen due to the reduction of the compressor flow rate and the turbine capacity, caused by a variation in the HPT nozzle throat area (geometric blockage due to the thickness of the deposited layer and the aerodynamic blockage due to the increased roughness, and in turn boundary layer). Several methods to quantify particle sticking have been proposed in literature so far, and the experimental data used for their validation vary in a wide range of materials and conditions. The experimental analyses have been supported by (and have given inspiration to) increasingly realistic mathematical models. Experimental tests have been carried out on (i) a full scale gas turbine unit, (ii) wind tunnel testing or hot gas facilities using stationary cascades, able to reproduce the same conditions of gas turbine nozzle operation and finally, (iii) wind tunnel testing or hot gas facilities using a coupon as the target. In this review, the whole variety of experimental tests performed is gathered and classified according to composition, size, temperature and particle impact velocity. Using particle viscosity and sticking prediction models, over seventy (70) tests are compared with each other and with the model previsions providing a useful starting point for a comprehensive critical analysis. Due to the variety of test conditions, the related results are difficult to be pieced together due to differences in particle material and properties. The historical data of particle deposition obtained over thirty (30) years are classified using particle kinetic energy and the ratio between particle temperature and its softening temperature. Qualitative thresholds for the distinction between particle deposition, surface erosion and particle break-up, based on particle properties and impact conditions, are identified. The outcome of this paper can be used for further development of sticking models or as a starting point for new insight into the problem.

The current trend in turbomachinery is pushing forward more and more efficient machines, increasing speeds, reducing components mass and improving their vibrational behaviour. Structural topology optimization is a challenging and promising approach to satisfy all these demands, with a very remarkable economic impact. This approach enables the creation of structures characterized by complex three-dimensional geometries, which are usually difficult or impossible to be produced using traditional manufacturing processes. However, thanks to innovative technologies, as new additive manufacturing techniques, it is now possible to effectively exploit topology optimization to develop innovative components. The aim of this work is to demonstrate the applicability of structural topology optimization techniques in turbomachinery, to improve the dynamic performances and vibrational behaviour of critical components. A 3D mock-up blade geometry based on T106 profile has been designed to reproduce a typical rotor blade in design conditions. The blade has been mounted on a rough disk model, to obtain a rotor blisk in order to ensure a wide design space for the optimization. The optimization has been carried out by applying mean and fluctuating loads coming from a 3D unsteady computation of 1.5stage (stator-rotor-stator) together with the centrifugal stresses. The unsteady loads acting on the rotor skin are due to the wake of the upstream stator and the potential field generated by the downstream stator. A new concept design for the blisk has been developed and the optimized geometry has been compared to the original one to highlight the improvements in terms of mass reduction and improved dynamic behaviour. This paper will confirm the suitability of this approach to turbomachinery components and a prototype of optimized geometry will be ready to be manufactured through innovative additive manufacturing techniques for high resistance alloys.

Components with best surface integrities — without any white layer or mechanically damaged rim zones — and regardless of mechanical properties of the material can be produced with electrochemical machining (ECM). Wire ECM combines the outstanding advantages of the ECM working principle with the kinematic of wire electrical discharge machining (wire EDM). A task-specific tool design — such as for classical sinking ECM — is not required for cutting 2.5-dimensional geometries like fir tree slots in turbine discs. Established manufacturing technologies for fir trees are broaching and recently wire EDM. However, high requirements on surface integrity of these turbine parts afford cost intensive finishing processes. Thus, wire ECM could be an attractive cost efficient alternative manufacturing process. Due to an insufficient flushing concept, a process development of wire ECM for fir tree slots with relevant disc heights failed in the past. A combination of an axial flushing with rotating structured electrode provides a promising approach to meet these challenges. Therefore, this paper presents investigations of cutting Inconel 718 direct-aged using the new flushing principle. Workpieces with heights up to 40 mm are machined, which is comparable to a low pressure turbine disc height. Reachable feed rates respectively cutting rates are presented. Due to the process specific edge rounding, the formation of the leading edges is examined. Finally, the evolving surface integrities with the help of surface roughness, cross sections and residual stresses are investigated and discussed.

Compressors are crucial to the efficient operation of a gas turbine; however, fouling, caused by adherence of particles to surfaces, can have a negative effect on compressor performance. In this study, a NASA Stage 35 single-stage axial-flow compressor was employed as the model for numerical simulation using the ANSYS CFX software of the effects caused by mild and severe fouling under salt fog scale. To measure these effects, two distinct models were used. For mild fouling, the simulated stator blade surface roughness was altered nonuniformly; for severe fouling, the simulated stator blade thick-ness was altered. Results indicated that surface roughness caused by mild fouling only has a small effect on compressor performance and no effect on the stable working range. However, changes in the blade thickness as a result of severe fouling have a large effect on compressor performance and a clear effect on the stability of the compressor’s working range. The fouling causes an increase in the boundary layer at the trailing edge of the suction side of the blade thereby increasing the loss of flow; fouling effect after emergence angle, wide design value, and increasing blade of circumferential stress.

The Marine Corps Systems Command is evaluating alternative propulsion systems to achieve high water speed for the future USMC Amphibious Combat Vehicle (ACV). A gas turbine engine is one of the propulsion systems evaluated. The primary risk of operating a gas turbine engine in the ACV is power loss due to the ingestion of marine contaminants such as saltwater mist in the air, saltwater spray generated from the vehicle operation, green water wash caused by the operation of the vehicle in the surf zone or in rough seas, and hard particles such as sand present in the marine environment. The Auxiliary Ships and New Acquisition Support Branch (Code 425) of the Naval Surface Warfare Center, Philadelphia Division conducted a study to assist the Marine Corps Systems Command in assessing the feasibility of using a gas turbine engine as a propulsion system on future USMC ACVs. The study was focused on developing and testing a gas turbine intake solution for the ACV that can remove saltwater from the intake airstream of the notional 3,000 horsepower ACV engine. Code 425 developed a two-part solution for the intake of the ACV. The first part of the solution is the Combustion Air Protection Shroud (CAPS) located at the entrance of the engine intake and designed to protect the ACV engine from green water wash by elevating the intake above the ACV deck. The second part of the solution is a gas turbine intake filtration system located downstream of the intake shroud and designed to remove marine contaminants that enter the intake shroud. A reduced-scale model of the CAPS was designed by Code 425 in conjunction with Gibbs & Cox and tested at the Davidson Laboratory High Speed Test Basin at the Stevens Institute of Technology to determine the optimum extension height of the CAPS to protect the engine intake. This paper covers the design and testing of the CAPS. The results showed that a 2.67 ft. tall CAPS with selectively closeable air intake louvers is sufficient to keep out saltwater from the ACV gas turbine engine.

Large eddy simulations of a linear low-pressure turbine cascade with the T106A profile and different surface roughness patches were carried out. The aim was to investigate the effects on the laminar and turbulent boundary layer on the blade suction surface. Two different approaches were used to represent the roughness patches. Firstly, a forcing model, reducing the computational costs compared to fully resolved roughness surfaces, was incorporated. Secondly, an immersed boundary method representing an as-cast roughness surface was used, for a more detailed analysis of flow mechanisms over roughness. It was found that the roughness model was able to induce boundary layer transition and alter the turbulent boundary layer, with the results in line with findings in the literature. The instantaneous flow data at different time instants of the as-cast roughness case showed the development of streaks due to distinct roughness peaks, resulting in highly uneven transition positions across the spanwise direction.

This paper experimentally addresses the impact of surface roughness on losses and secondary flow in a Turbine Rear Structure (TRS). Experiments were performed in the Chalmers LPT-OGV facility, at an engine representative Reynolds number with a realistic shrouded rotating low-pressure turbine (LPT). Outlet Guide Vanes (OGV) were manufactured to achieve three different surface roughnesses tested at two Reynolds numbers, Re = 235000 and Re = 465000. The experiments were performed at on-design inlet swirl conditions. The inlet and outlet flow of the TRS were measured in 2D planes with a 5-hole probe and 7-hole probe accordingly. The static pressure distributions on the OGVs were measured and boundary layer studies were performed at the OGV midspan on the suction side with a time-resolved total pressure probe. Turbulence decay was measured within the TRS with a single hot-wire. The results showed a surprisingly significant increase in the losses for the high level of surface roughness (25–30 Ra) of the OGVs and Re = 465000. The increased losses were primary revealed as a result of the flow separation on the OGV suction side near the hub. The loss increase was seen but was less substantial for the intermediate roughness case (4–8 Ra). Experimental results presented in this work provide support for the further development of more advanced TRS and data for the validation of new CFD prediction methods for TRS.

High-turbulence flow conditions in gas turbine combustors create a rough environment for film cooling flows. Fundamental knowledge of the interaction between main flow fluctuations and cooling jet is essential to develop more efficient cooling geometries. To gain detailed insight into the turbulent flow structures, high-speed PIV measurements have been carried out in a closed loop, thermal wind tunnel facility. Velocity fields with a temporal resolution of 6 kHz and a vector spacing of 0.5 mm have been extracted from the recorded data. Measurements have been performed for a density ratio of DR = 1.2 and momentum ratios of I = 2, I = 8 and I = 15. A Makita style active turbulence grid has been used to create a main flow turbulence intensity of Tu = 18.5% and Tu = 21.5%. Mean flow fields and dominant flow structures that are present in the time-resolved velocity fields are analyzed and discussed. Finally, results for high-turbulence main flow conditions are compared with results from low-turbulence (Tu = 1%) thermographic PIV data, carried out in previous measurements. It can be observed that low-momentum jets are highly fragile. Both small and large vortices affect the jet trajectory. Depending on the current flow situation, the jet lifts off the surface or gets pushed towards the wall. The same is true for high momentum ratios and large eddies. On the other hand, small eddies cannot significantly affect high-momentum jets. The vortex gets deflected by the jet and its intensity is reduced while interacting with the high-turbulence shear layer fluctuations. All in all, turbulence can have a positive effect on the cooling film by steering the jet trajectory towards the surface. On the other hand, however, vortices can enhance the jet detachment and thus reduce the cooling efficiency of the film.

The degree of complexity in internal cooling designs is tied to the capabilities of the manufacturing process. Additive manufacturing grants designers increased freedom while offering adequate reproducibility of micro-sized, unconventional features that can be used to cool the skin of gas turbine components. One such desirable feature can be sourced from nature; a common characteristic of natural transport systems is a network of communicating channels. In an effort to create an engineered design that utilizes the benefits of those natural systems, the current study presents wavy microchannels that were connected using branches. Two different wavelength baseline configurations were designed, then each were numerically optimized using a commercial adjoint-based method. Three objective functions were posed to (1) minimize pressure loss, (2) maximize heat transfer, and (3) maximize the ratio of heat transfer to pressure loss. All baseline and optimized microchannels were manufactured using Laser Powder Bed Fusion for experimental investigation; pressure loss and heat transfer data were collected over a range of Reynolds numbers. The additive manufacturing process reproduced the desired optimized geometries faithfully. Surface roughness, however, strongly influenced the experimental results; successful replication of the intended flow and heat transfer performance was tied to the optimized design intent. Even still, certain test coupons yielded performances that correlated well with the simulation results.

High performance film cooling holes with complicated geometries have been regarded as impractical up to now because of manufacturability issues. However, recent advances in additive manufacturing (AM) technology have opened up new doors. Investigating characteristics of film holes built with AM, and finding the optimum shape considering these characteristics are now required to confirm their practical utility. In this paper, the performance of a high-efficiency film hole is numerically investigated. In-hole roughness and blade surface roughness are examined assuming an AM process, and contorted hole shape caused by partial blockage is also considered. A robust hole shape is obtained considering these uncertainties, utilizing a reference hole shape made by combining three cylindrical holes. Hole diameter, injection angle, and two angles for defining the two auxiliary holes are used as design variables to be optimized. For flow field and thermal analysis with roughness, compressible steady Reynolds averaged Navier-Stokes equations with a sand-grain roughness model are used. For the probabilistic assessment of each hole shape, Monte Carlo Simulations with the Kriging surrogate model is used, along with efficient global optimization (EGO) and a genetic algorithm. As a result, a high performance yet robust film cooling hole shape is obtained.

Heat transfer in annuli has been widely investigated over the years. Annular channels are found in multiple industry applications, ranging from electronic equipment to nuclear reactor cooling. In order to enhance heat transfer, internal passages are augmented with roughness elements such as dimples, pimples, ribs, fins, and other features. However, the use of impingement cooling mechanism has not been utilized for such annular channels. An experimental study of impingement cooling within an annulus with the outer surface of the annulus being heated uniformly is conducted. Multiple flows as a function of Jet Reynolds number in the range of 16,000 to 46,000 are fundamentally examined. An impingement sleeve with an annular ratio of 0.92 was experimentally tested and compared to RANS numerical simulation results. Temperature Sensitive Paint (TSP) was utilized to experimentally obtain the local Heat Transfer distribution along the annular wall and close channel cap, which is normalized by the baseline annulus heat transfer distribution. Heat transfer results for the annular surface is compared with available correlations. Pressure drop across the facility is characterized at the inlet and exit of the experimental set up.

Gas turbine components subjected to high temperatures can benefit from improved designs enabled by metal additive manufacturing with nickel alloys. Previous studies have shown that the impact on fluid flow and heat transfer resulting from surface roughness of additively manufactured parts is significant; these impacts must be understood to design turbine components successfully for additive manufacturing. This study improves understanding of these impacts by examining the discharge coefficient and the effect of the coolant delivery direction on the performance of additively manufactured shaped film cooling holes. To accomplish this, five test coupons containing a row of baseline shaped film cooling holes were made from a high temperature nickel alloy using a laser powder bed fusion process. Flow and pressure drop measurements across the holes were collected to determine the discharge coefficient from the film cooling holes. Temperature measurements were collected to assess the overall effectiveness of the coupon surface as well as the cooling enhancement due to film cooling. The Biot number of the coupon wall was matched to a value one might find in a turbine engine to ensure this data is relevant. It was discovered that the flow experienced greater aerodynamic losses in film cooling holes with greater relative roughness, which resulted in a decreased discharge coefficient. The effectiveness measurements showed that the film cooling performance is better when coolant is fed in a co-flow configuration compared to a counter-flow configuration.

Over recent decades, the variability and high costs of the traditional gas turbine fuels (e.g. natural gas), have pushed operators to consider low-grade fuels for running heavy-duty frames. Synfuels, obtained from coal, petroleum or biomass gasification, could represent valid alternatives in this sense. Although these alternatives match the reduction of costs and, in the case of biomass sources, would potentially provide a CO 2 emission benefit (reduction of the CO 2 capture and sequestration costs), these low-grade fuels have a higher content of contaminants. Synfuels are filtered before the combustor stage, but the contaminants are not removed completely. This fact leads to a considerable amount of deposition on the nozzle vanes due to the high temperature value. In addition to this, the continuous demand for increasing gas turbine efficiency, determines a higher combustor outlet temperature. Current advanced gas turbine engines operate at a turbine inlet temperature of (1400–1500) °C which is high enough to melt a high proportion of the contaminants introduced by low-grade fuels. Particle deposition can increase surface roughness, modify the airfoil shape and clog the coolant passages. At the same time, land based power units experience compressor fouling, due to the air contaminants able to pass through the filtration barriers. Hot sections and compressor fouling work together to determine performance degradation. This paper proposes an analysis of the contaminant deposition on hot gas turbine sections based on machine nameplate data. Hot section and compressor fouling are estimated using a fouling susceptibility criterion. The combination of gas turbine net power, efficiency and turbine inlet temperature (TIT) with different types of synfuel contaminants highlights how each gas turbine is subjected to particle deposition. The simulation of particle deposition on one hundred (100) gas turbines ranging from 1.2 MW to 420 MW was conducted following the fouling susceptibility criterion. Using a simplified particle deposition calculation based on TIT and contaminant viscosity estimation, the analysis shows how the correlation between type of contaminant and gas turbine performance plays a key role. The results allow the choice of the best heavy-duty frame as a function of the fuel. Low-efficiency frames (characterized by lower values of TIT) show the best compromise in order to reduce the effects of particle deposition in the presence of high-temperature melting contaminants. A high-efficiency frame is suitable when the contaminants are characterized by a low-melting point thanks to their lower fuel consumption.

The average roughness height Ra was used for characterizing surface roughness in many past investigations. In order to fully describe heat transfer and fluid flow over a rough surface, statistical parameters, however, are not sufficient. There is a lack of data in the literature in terms of three-dimensional real turbine roughness and the distribution of roughness parameters on the turbine airfoils. This study closes this gap and provides detailed surface roughness data of high-pressure turbine vanes that have been in service in land-based gas turbines and aero engines. Furthermore, samples of thermal barrier coatings from two different application techniques are investigated. High-resolution 3D profilometer and optical measurements are performed at various locations on both the pressure and suction side. Key parameters such as roughness shape and density are quantified. Equivalent sand grain roughness height (ks) values are obtained and related to the fluid flow by calculating the corresponding dimensionless roughness heights. The results show that roughness parameters are spread over a wide range. No correlation with position on the vane surface is found. The distribution of the roughness heights deviates from a Gaussian distribution. Roughness on the suction side is generally smaller than on the pressure side. Dimensionless roughness heights are found to be mainly in the transitional regime and not in the fully rough regime as expected. The data presented in this paper is of great importance to the design of test surfaces for laboratory experiments and as input for numerical models.