Reversible axial fans are widely used in industrial and tunnel ventilation systems, and a lot of research effort is spent in the design process of the blades shape and blades profile. The target is to achieve reasonable performances in both flow directions, but those are still below the levels of the corresponding non-reversible geometries. In this paper, an alternative design solution for reversible axial fan is presented by adopting flexible blades instead of the rigid ones. Such design, inspired by the boat sails, could allow the blade to change its shape by passively adapting to the flow field, from a symmetrical blade profile to a not symmetric one, and thus adapting the curvature to the flow condition. In the paper, a series of alternative materials and material distributions are analysed and compared. The analysis is conducted by performing Fluid-Structure Interaction simulations using stabilized Finite Elements formulations for both the fluid and the structure dynamics. Simulations are performed using the in-house built software FEMpar, which implements the Residual Based Variational MultiScale to model the Navier-Stokes equation, the Total Lagrangian formulation for the non-linear elastic solid and the Solid Extension Moving Mesh Technique to move the fluid mesh.

This article presents high-fidelity experimental traverse measurements downstream of an annular cascade of transonic nozzle guide vanes (NGVs) from a high-pressure (HP) turbine stage. The components are heavily cooled real engine components from a modern civil gas turbine engine, operated at scaled engine conditions. Tests were conducted in the high technology readiness level (TRL) Engine Component Aerothermal (ECAT) facility at the University of Oxford. High-resolution full-area traverse measurements of local kinetic energy (KE) loss coefficient are presented in several axial planes. In particular, we present circumferential loss coefficient profiles at several radial heights, full-area traverses at three axial planes, and fully mixed-out loss calculations. The analysis of these data gives insight into particular loss structures, overall aerodynamic performance, and wake mixing rates. The effect of exit Mach number on performance is also considered. The data address a gap in the literature for the detailed analysis of traverse measurements downstream of HP NGV engine components. Experimental data are compared with steady and unsteady Reynolds-averaged Navier–Stokes (RANS) simulations, allowing benchmarking of typical computational fluid dynamics (CFD) methods for absolute loss prediction of cooled components. There is relatively limited aerodynamic performance data in the literature for heavily cooled NGVs, and this study represents one of the most comprehensive of its type.

This paper presents a comparative experimental and numerical study of the heat transfer and pressure loss in a swirl multi-pass channel with tangential jet slots, and another baseline multi-pass channel with 180-deg U-bends as comparison baseline has also been investigated. Transient liquid crystal thermography is used to obtain the detailed heat transfer distribution on the internal surfaces of the multi-pass serpentine channels. The heat transfer patterns in the swirl multi-pass channel are quite different from that of the baseline multi-pass channel. Compared with the baseline multi-pass channel, the experimental globally averaged Nusselt number ratios of the last two passes in the swirl multi-pass channel can be increased by up to 82.9%, 104.8% and 124.6% for the Reynolds numbers 20,000, 40,000 and 60,000, respectively. The high and circumferentially uniform heat transfer is mainly due to the large-scale swirling flow induced by the tangential slots. More detailly, the large-scale swirling flow impinges onto the surface and further induces high tangential velocity near the wall, which destroys the boundary layer flow and thus improves the heat transfer rates at the wall. However, the notable pressure loss of the swirl multi-pass channel should be further controlled reasonably, which is about 5.4 times that of the baseline multi-pass channel. As supplements to the experiments, three-dimensional numerical computations provide more insights into the turbulent flow structure in the two kinds of multi-pass serpentine channels.

A wide variety of pin-fins have been used to enhance heat transfer in internal cooling channels. However, due to their large blockage in the flow direction, they result in an undesirable high pressure drop. This experimental study aims to reduce pressure drop while increasing the heat transfer surface area by utilizing strip-fins in converging internal cooling channels. The channel is designed with a trapezoidal cross-section, converges in both transverse and longitudinal directions, and is also skewed β=120° with respect to the direction of rotation in order to model a trailing edge cooling channel. Only the leading and trailing surfaces of the channel are instrumented, and each surface is divided into eighteen isolated copper plates to measure the regionally averaged heat transfer coefficient. Utilizing pressure taps at the inlet and outlet of the channel, the pressure drop is obtained. Three staggered arrays of strip-fins are investigated: one full height configuration and two partial fin height arrangements (Sz=2mm and 1mm). In all cases, the strip fins are 2mm wide (W) and 10mm long (Lf ) in the flow direction. The fins are spaced such that Sy/Lf = 1 in the streamwise direction. However, due to the convergence, the spanwise spacing, Sx/W, was varied from 8 to 6.2 along the channel. The rotation number of the channel varied up to 0.21 by ranging the inlet Reynolds number from 10,000 to 40,000 and rotation speed from 0 to 300rpm. It is found that

In the recent years Additive Manufacturing (AM) methods are getting more and more attractive and feasible for the realization of components and subcomponents of gas turbines. They are receiving much attention since, on one hand, the manufacturing of complex 3D geometries is allowed and, on the other, manufacturing and delivery times can be cut down. At the current state of the art, to the authors' knowledge only few applications have yet been commercialized relatively to cooling holes, due to the intrinsic difficulties associated with such a critical feature. Lately, Baker Hughes is studying the possibility to manufacture film-cooling holes via the DMLM technology in order to exploit the flexibility of such innovative manufacturing method and hence eliminate additional processes and lead time. From the open literature it is known that additively manufactured holes can have a more irregular shape and higher roughness than traditional ones, which may lead not only to a reduction in coolant flow but more importantly to a decay of the film-cooling adiabatic effectiveness. For this reason, a test campaign has been conducted in collaboration with the University of Florence (Italy) with the objective of characterizing the performance (minimum passage diameter, flow check and adiabatic effectiveness) of AM vs traditional cylindrical holes on simple-geometry coupons built upon different construction angles. Results were then analyzed in order to fully compare the performance of AM vs traditional film-cooling holes at different operating regimes.

Blades manufactured through flank and point milling will likely exhibit geometric variability. Gauging the aerodynamic repercussions of such variability, prior to manufacturing a component, is challenging enough, let alone trying to predict what the amplified impact of any in-service degradation will be. While rules of thumb that govern the tolerance band can be devised based on expected boundary layer characteristics at known regions and levels of degradation, it remains a challenge to translate these insights into quantitative bounds for manufacturing. In this work, we tackle this challenge by leveraging ideas from dimension reduction to construct low-dimensional representations of aerodynamic performance metrics. These low-dimensional models can identify a subspace which contains designs that are invariant in performance -- the inactive subspace. By sampling within this subspace, we design techniques for drafting manufacturing tolerances and for quantifying whether a scanned component should be used or scrapped. We introduce the blade envelope as a computational manufacturing guide for a blade. In this paper, the first of two parts, we discuss its underlying concept and detail its computational methodology, assuming one is interested only in the single objective of ensuring that the loss of all manufactured blades remains constant. To demonstrate the utility of our ideas we devise a series of computational experiments with the Von Karman Institute's LS89 turbine blade.

Blade envelopes offer a set of data-driven tolerance guidelines for manufactured components based on aerodynamic analysis. In Part I of this two-part paper, a workflow for the formulation of blade envelopes is described and demonstrated. In Part II, this workflow is extended to accommodate multiple objectives. This allows engineers to prescribe manufacturing guidelines that take into account multiple performance criteria. The quality of a manufactured blade can be correlated with features derived from the distribution of primal flow quantities over the surface. We show that these distributions can be accounted for in the blade envelope using vector-valued models derived from discrete surface flow measurements. Our methods result in a set of variables that allows flexible and independent control over multiple flow characteristics and performance metrics, similar in spirit to inverse design methods. The augmentations to the blade envelope workflow presented in this paper are demonstrated on the LS89 turbine blade, focusing on the control of loss, mass flow and the isentropic Mach number distribution. Finally, we demonstrate how blade envelopes can be used to visualize invariant designs by producing a 3D render of the envelope using 3D modelling software.

There is a need to improve predictions of losses resulting from large eddy simulations (LES) of low-pressure turbines (LPT) in gas turbines. This may be done by assessing the accuracy of predictions against validation data and understanding the source of any inaccuracies. LES is a promising approach for capturing the laminar/turbulent transition process in a LPT. In previous studies, the authors utilized LES to model the flow field over a Variable Speed Power Turbine (VSPT) blade and successfully captured characteristic features of separation/reattachment and transition on the suction side at both the cruise (positive incidence) and take-off conditions (negative incidence) and as well, simulated the effect of freestream turbulence (FST) on those phenomena. The predicted pressure loading profiles agreed well with the experimental data for both a high and a low FST case at a Reynolds number of Re ex = 220,000. In this paper, we present wake profiles resulting from computations for a range of FST values. Although the predicted wake profiles for the lowest FST case (Tu = 0.5%) matched the experimental data, at higher FST (Tu = 10-15%,) the wake was wider than the experimentally measured wake and for both cases were displaced laterally when compared to the experimental measurements. In our investigation of the causes of the said discrepancies we have identified important effects which could strongly influence the predicted wake profile. Predicted losses were improved by assuring the validity of the flow solution.

Every supersonic fan or compressor blade row has a streamtube, the “sonic streamtube,” which operates with a blade relative inlet Mach number of one. A key parameter in the design of the “sonic streamtube” is the area ratio between the blade throat area and the upstream passage area, A throat /A inlet . In this article, it is shown that one unique value exists for this area ratio. If the area ratio differs, even slightly, from this unique value, then the blade either chokes or has its suction surface boundary layer separated due to a strong shock. Therefore, it is surprising that in practice designers have relatively little problem designing blade sections with an inlet relative Mach number close to unity. This article shows that this occurs due to a physical mechanism known as “transonic relief.” If a designer makes a mistake and designs a blade with a “sonic streamtube,” which has the wrong area ratio, then “transonic relief” will self-adjust the spanwise streamtube height automatically moving it toward the unique optimal area ratio, correcting for the designer’s error. Furthermore, as the blade incidence changes, the spanwise streamtube height self-adjusts, moving the area ratio toward its unique optimal value, effectively controlling the blade’s incidence range. Without “transonic relief,” supersonic and transonic fan and compressor design would be impossible. This article develops a simple model that allows “transonic relief” to be decoupled from other mechanisms and to be systematically studied. The physical mechanism on which it is based is thus determined and a universal relationship for core compressor preliminary blade design is presented. Finally, its implications in relaxing manufacturing tolerances and in the design of future distortion tolerant blades are discussed.

A common objective in the analysis of turbomachinery components (nozzle guide vanes or rotor blades, for example) is to calculate performance parameters, such as total pressure or kinetic energy loss coefficients, from measurements in a non-uniform flow-field. These performance parameters can be represented in a range of ways. For example: line-averages used to compare performance between different radial sections of a 3D component; plane-averages used to assess flow (perhaps loss coefficient) development between different axial planes; and fully mixed-out values used to determine the total loss associated with a component. In this paper, we compare a range of methods for calculating aerodynamic performance parameters including plane-average methods with different weighting schemes and several mixed-out methods. We analyse the sensitivities of the different methods to the axial location of the measurement plane, the radial averaging range, and the exit Mach number. We use high-fidelity experimental data taken in several axial planes downstream of a cascade of engine parts: high pressure (HP) turbine nozzle guide vanes (NGVs) operating at transonic Mach number. The experimental data is complemented by CFD. We discuss the underlying physical mechanisms which give rise to the observed sensitivities. The objective is to provide guidance on the accuracy of each method in a relevant, practical application.

In the current study, the tandem blade technology is applied to an supersonic throughflow fan (STFF) tandem cascade for the first time and a 2D STFF tandem cascade is preliminarily designed. Through the modification design of the tandem airfoils and their configuration (axial overlap (AO) and percent pitch (PP)), the coefficients of total pressure loss and loading are reduced by 4% and 8.58%, respectively. Furthermore, the impact of tandem configurations on the performance is parametrically investigated by numerical simulations. The results indicate that compared with AO, the performance under design incidence is more sensitive to PP except for the cases with PP exceeding a threshold value (1.15). PP dominates the loss and load by controlling the evolution of the FB wake and the shock structure of FB and the rear one of the tandem blade (RB), while AO mainly adjusts the entire shock system structure through the change of virtual shape, resulting in the variation in load distribution between FB and RB. It is worth noting that the overall loading and the total loss remain unchanged with increasing AO except for the tandem configurations (PP = 1.05, AO ≤ −0.01), which make the flow structure in the gap region undergo a fundamental change. With the optimal tandem configuration (PP = 1.05, AO = −0.01) and the modified tandem blades (the ratios of chord length and camber for FB over RB are 0.67 and 0.5, respectively), the total pressure loss coefficient is further reduced by 19.7% in comparison with the preliminary tandem design.

The manufacturing geometrical variability is a source of uncertainty, which cannot be avoided in the realization of machinery components. Deviations of a part geometry from its nominal design are inevitably present due to the manufacturing process. In the case of the aeroelastic forced response problem within axial compressors, these uncertainties may affect the vibration characteristics. For this reason, the impact of geometrical uncertainties due to the manufacturing process onto the modal forcing of axial compressor blades is investigated in this study. The research focuses on the vibrational behavior of an axial compressor rotor blisk. In particular, the amplitude of the forces acting as a source of excitation on the vibrating blades is studied. The geometrical variability of the upstream stator is investigated as input uncertainty. The variability is modeled starting from a series of optical surface scans. A stochastic model is created to represent the measured manufacturing geometrical deviations from the nominal model. A data reduction methodology is proposed in order to represent the uncertainty with a minimal set of variables. The manufacturing geometrical variability model allows to represent the input uncertainty and probabilistically evaluate its impact on the aeroelastic problem. An uncertainty quantification is performed in order to evaluate the resulting variability on the modal forcing acting on the vibrating rotor blades. Of particular interest is the possible rise of low engine orders due to the mistuned flow field along the annulus. A reconstruction algorithm allows the representation of the variability during one rotor revolution. The uncertainty on low harmonics of the modal rotor forcing can be therefore identified and quantified.

We explain the advantages of using big-data methods for evaluating numerical optimization. The investigation focuses on the performance potential of three-dimensional return channel vanes under realistic manufacturing constraints. Based on an analysis of an optimization database, this paper presents a systematic approach for analysis and design guidelines for three-dimensional return channel vanes. To this end, a validated numerical setup was developed on the basis of experimental investigations, followed by a numerical optimization using a genetic algorithm and an artificial neural network. The optimization database was analyzed with a dimension reduction method called t-stochastic neighbor embedding. This method enabled linking geometric design features with physical correlations and, finally, with the objective functions of the optimization. With the help of the detected correlations within the database, it has been possible to work out a method for deciding on the selection of a design on the Pareto front and to draw new relevant conclusions. The systematic use of big-data methods proposed enables a more penetrating insight to be gained into numerical optimization, which is more general and relevant than those gained by simply comparing a single optimized design with a reference design. An analysis of the Pareto front reveals that 0.6% efficiency can be exchanged for 20% more homogeneous outflow.

Variable inlet guide vanes (VIGV) are the main control element to adjust the flow rate of industrial centrifugal compressors by customized pre-swirl in the inlet plane of the impeller. The efficient working range of VIGVs is however restricted due to open flow separation occurring at critical stagger angles. In order to overcome the narrow limitations of current blade geometries and to enhance the operating range of the compressor, split blades consisting of a separate front and tail blade segment proved to be particularly promising in previous linear-cascade measurements. Each blade segment is thereby individually staggered. This enables a gradual flow deflection along the chord length. Secondary flow losses, however, were not considered in the previous investigations with linear cascades. To highlight the potential of the split blade concept under more application-oriented conditions including all relevant flow effects, highly resolved field measurements were conducted in the wake of annular VIGVs. Four different blade configurations, a customary reference case and three variations of the split blade with full, partial and missing sealing in the gap between the segments were assessed using five-hole probe measurements. By investigating a wide range of stagger angles, the coverage of the full low-loss working range of the VIGV could be ensured. Especially, the fully sealed split blade configuration proved its capacity to extend the efficient operational range significantly.

Nickel-base superalloys are extensively used in the production of gas turbine hot-section components as they offer exceptional creep strength and superior fatigue resistance at high temperatures. Such improved properties are due to the presence of precipitate-strengthening phases such as Ni 3 Ti or Ni 3 Al (γ′ phases) in the normally face-centered cubic (FCC) structure of the solidified nickel. Although this second phase is the main reason for the improvements in properties, the presence of such phases also results in increased processing difficulties as these alloys are prone to crack formation. In this work, specimens of IN738LC are fabricated on a Coherent Creator laser powder bed fusion (L-PBF) additive manufacturing (AM) equipment. Optical microscopy (OM), scanning electron microscopy (SEM), electron backscatter diffraction (EBSD), and X-ray diffraction (XRD) are carried out to characterize the deposit region. Metallurgical continuity is achieved in the entire deposit region and the specimens do not show any warpage. However, the specimens show voids (e.g., pores and cracks) in the deposit region. The results show that the percentage void area decreases along the build height direction. The deposited IN738LC shows polycrystalline grains in the entire deposit region as confirmed by XRD and EBSD. The grain size also shows variations along the build direction. In summary, the results open opportunities for academic researchers and small-scale businesses in fabricating high-γ′ nickel-base superalloys on a desktop laser powder bed fusion AM equipment.

Impingement cooling is one of the powerful cooling methods in high-temperature devices. For the gas turbine applications, impingement cooling is commonly applied in the transition piece of a combustor and in the leading edge, suction and pressure sides of a turbine blade/vane. In the suction side and pressure side, impingement cooling is applied as a form of an array jet. However, due to the small gap between the jet hole and target surface, the wall jet faces a crossflow inside of the gap. This crossflow has an adverse effect on the jets and deteriorates the heat/mass transfer performance. Therefore, several studies have been conducted to minimize the crossflow effect. The present study also investigated the effect of crossflow reduction in the gap by having a castellated hole plate. The heat/mass transfer was measured using the naphthalene sublimation method. Heat/mass transfer data are compared among three different cases. One is the baseline case which is simple array impinging jets. Others are the castellated cases with and without rib structures on the target wall. Jet-to-jet spacing(s/d) and jet-to-target spacing(z/d) are selected as geometrical variables. Also, the experiments were conducted for the Reynolds numbers (based on jet hole diameter) of 5,000, 15,000, and 30,000. The baseline case was named as B case, and the castellated case without rib structure as C case and with rib structure as CR case. Both C and CR cases showed the crossflow reduction effect and resulted high and similar Nusselt number values.

Part II describes the experimental assessment and the application of the ideas in Part I concerning the mechanisms that determine the role of blade surface waviness on laminar-turbulent transition and their consequent effect on civil aircraft fan performance. A natural transition wind tunnel was designed and constructed to characterize the impact of surface waviness on transition, using both hotwire anemometry and infrared thermography. The experimental results support the new hypothesis presented in Part I , concerning the way in which blade surface waviness affects fan performance through motion of the transition onset location due to interaction between surface waviness and Tollmien–Schlichting (TS) boundary layer instability. In particular, the theoretical amplification of the TS waves, and the corresponding transition onset location movement due to surface waviness, was borne out over a range of variations in Reynolds number, nondimensional surface wavelength, nondimensional surface wave height, and location of surface wave initiation, relevant to composite fan blade parameters. Further, the increase of receptivity coefficient, and thus, the initial amplitude of disturbances due to geometric resonance between surface wavelength and TS wavelength was also confirmed by the experiments. Surface waviness was estimated, in some cases, to result in a nearly 1% decrease in fan efficiency compared to a nonwavy blade. Suggestions are given for mitigation of the effects of waviness, including the idea of blade curvature rescheduling as a method to delay transition and thus decrease loss.

The need for high performances is pushing the complexity of mechanical design at very high levels, especially for turbomachinery components. In this field, structural topology optimization methods together with additive manufacturing techniques for high-resistant alloys are considered very promising tools, but their potentialities have not been deeply investigated yet for critical rotating components like new-generation turbine blades. In this context, this research work proposes a methodology for the design, the optimization, and the additive manufacturing of extremely stressed turbomachinery components like turbine blade rows. The presented procedure pays particular attention to important aspects as fluid–structure interactions (forced response and flutter phenomena) and fatigue behavior of materials, going beyond the standard structural optimization approaches. The new design strategy enables a substantial reduction of the component mass, limiting the maximum stress and improving the vibrational behavior of the system in terms of eigenfrequencies, modal shapes, and fatigue life. Furthermore, the numerical procedure shows robustness and efficiency, making the proposed methodology an appropriate tool for rapid design and prototyping and for reducing the design costs and the typical time to market of this type of mechanical elements. The procedure has been applied to a low-pressure turbine rotor to improve the aeromechanical characteristics while keeping the aerodynamic performance. From the original geometry, mode-shapes, forcing functions (due to rotor/stator interactions), and aerodynamic damping have been numerically evaluated and used as input data for the subsequent topological optimization. Finally, the optimized geometry has been numerically and experimentally verified to confirm the improved aeromechanical design. After the structural topology optimization, the final geometries provided by the procedure have been properly rendered to make them suitable for additive manufacturing. Some prototypes of the new optimized turbine blade have been manufactured from aluminum alloy to be tested mechanically and in terms of fatigue.

Transpiration cooling represents the pinnacle of turbine cooling and is characterized by an intrinsic material porosity which achieves high internal convective cooling, and full coverage cooling films on the external surface subjected to the hot gases. Quasi-transpiration systems, such as the double-wall effusion system discussed here, attempt to replicate the cooling effect of transpiration systems. The double-wall system is characterized by a large internal wetted area providing high internal convective cooling performance, with a highly porous external wall allowing the formation of a protective film over the external surface. This paper presents a low-order thermal model of a double-wall system designed to rapidly ascertain cooling performance based solely on the geometry, solid thermal conductivity, and approximate surface heat transfer coefficients. The performance of the model is initially validated using experimental data with heat transfer coefficients for the low-order model obtained from fully conjugate computational fluid dynamics (CFD) simulations. Following this, a more controlled CFD study is undertaken with both fully conjugate and fluid-only simulations performed on several double-wall geometries to ascertain both overall effectiveness and film effectiveness data. Data from these simulations are used as inputs to the low-order thermal model developed and the results compared. The low-order model successfully captures both the trends and absolute cooling effectiveness achieved by the various double-wall geometries. The model therefore provides an extremely powerful tool in which the cooling performance of double-wall geometries can be near instantaneously predicted during the initial design stage, potentially allowing geometry optimization to rapidly occur prior to more in-depth, costly, and time-consuming analyses of the systems being performed. This potential benefit is demonstrated via the implementation of the model with input boundary conditions obtained using empirical correlations.

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 (CRTs) 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 (HTRs) approaching unity, with matched non-dimensional design parameters, the stage efficiency and work output are identical for both types of machines. However, a CRT in the study is shown to have an efficiency advantage over a conventional machine of up to 0.35 percentage points for a HTR of 0.65. This is due to differences in absolute velocity producing different spanwise blade designs.