Criteria for assessing high cycle fatigue (HCF) capability is important for transitioning additive repair technologies to turbine engine applications. By studying the fatigue results of two laser directed energy deposition additive manufacturing repairs on airfoil representative Titanium 6Al-4V coupons, acceptable HCF results for repairs can be determined by observing the confidence in predicted values and the reliability of the empirical data compared to baseline results. The following document details the steps and philosophies behind fabricating coupons that capture the repair capability. Furthermore, regression analysis as well as investigations of dataset distribution, fractography, and microscopy are necessary for instilling confidence in the understanding of repair fatigue behavior. The findings in this work provide a decision gate with quantifiable metrics for advancing novel repair techniques.

To meet today’s and future market needs, large end-stage blades are obliged to fulfill high flexibility regarding the operational range and high efficiency goals while being prepared for daily start-stop cycles. The end-stage total efficiency can be maximized by enlarging the steam turbine exhaust area and thereby reducing the exhaust losses. Therefore, a new Low Pressure (LP) backend featuring an increased freestanding 41″ steel blade has been developed and is presented here, which is optimized for maximum efficiency over a wide range of operation conditions. To allow for such a large steel-blade to operate at 60Hz rotational speed and to meet the daily cycling demand, various aspects of the blade design were optimized. A new high strength blade steel was developed (Teuber [1]), which gives the designer freedom for aerodynamical optimizations, while keeping the mechanical utilization within the predefined, allowable limits. To maximize the cycling capability, a new fir tree root was developed which minimizes the static as well as the dynamic loading. To verify the success of the new fir-tree root design and to verify the natural frequencies for the relevant modes, an extensive validation measurement campaign was setup with a full-scale blade row in a spin-pit. Here, the airfoil, root and steeple of the end-stage blade were equipped with strain gauges. Additionally, the blade row was monitored using tip-timing sensors. The results of this validation measurement campaign are presented in this paper. They show a close agreement between the design calculations and the measured static strains and vibration responses in terms of natural frequencies as well as displacement and strain amplitudes. Additionally, a test turbine has been set-up featuring a direct scaling of the new LP backend with the new high strength steel and a pre-stage to simulate realistic operation conditions over the complete operation range. The blade performance was tested up to high mass-flows, condenser pressures of up to 300 mbar and at varying load points covering all potential load points from extreme part load to full load with minimal and maximal condenser pressure. Strain gauges as well as tip-timing are used to measure the vibration response of the end-stage blade during the measurement campaign. The results presented here show, that throughout the complete measurement campaign the blade experienced minimal excitation which led to vibration levels that allowed unrestricted operation in the complete, tested operation range. In summary this paper shows the main design features of a large full-speed freestanding end-stage blade and the validation measures that were performed to ensure that the design targets and the market requirements are fully met.

A full-size, full-speed, axial flow steam turbine test rig capable of measuring turbine thrust, and static pressures in the rotor-stator disk cavity was built and commissioned. The test rig was operated in a single-stage configuration for the test results first reported in Stasenko et al. [1], and now in this paper. The stage has stationary axial face seals radially inward of the airfoils, near the rotor disk rim. The face seals divide the rotor-stator cavity into inner and outer circumferential cavities, both of which were instrumented with static pressure probes on the stator radial wall. Axial thrust was measured with load cells in every thrust bearing pad. The test rig was operated over a range of three nominal stage pressure ratios (designated as LPR, MPR, and HPR), five nominal stage velocity ratios (0.25–0.6), and five admission fractions (0.38–0.88). This latest group of tests was conducted without rotor disk balance holes, which were mechanically plugged, and will be compared to the original block of tests with disk balance holes opened. In the upstream disk cavity, the two disk balance hole configurations shared many similar pressure characteristics: nearly uniform pressures in the inner cavity, circumferential pressure distributions in the outer cavity that corresponded with the direction of axial thrust, and radial pressure distributions in the outer cavity that were a direct function of rotor speed. General trends of thrust coefficients with the disk holes plugged were correlated to stage pressure ratio, stage velocity ratio, admission fraction, and leakage mass flow rate. Those trends were consistent with the first block of tests with open disk balance holes, although there was an offset toward more operating conditions with negative aggregate thrust coefficients. This suggests that the rotating disk induces a low-pressure gradient in the inner (upstream) cavity, and the opened disk balance holes tend to equalize the inner cavity static pressure toward the higher static pressure on the exit side of the disk. Additionally, thrust coefficients tended to become less negative (or more positive) with stage pressure ratio and with velocity ratio, but tended to become more negative with admission fraction. Significant thrust coefficient reductions were realized with the open disk balance hole configuration, and were determined to be consistently speed-dependent.

In this paper, a novel model is presented for reconstructing unsteady periodic fields of velocity vector and pressure scalar over an oscillating foil. This data-driven method based on convolutional neural network can be utilized to accomplish two objections: fields reconstruction from limited measurements and transient aerodynamic characteristics prediction. The verification results of an oscillating foil under low Reynolds number show that this method can accurately reconstruct all the fields only by limited pressure information at probes on the foil surface. The evaluation on aerodynamic characteristics prediction illustrates that our model outperforms four classical machine learning methods. Meanwhile, a well-trained CNN model can almost achieve real-time flow field prediction by leveraging the GPU acceleration. Finally, the exploration of the robustness for the CNN model is conducted on several aspects, including training size, probe layouts, probe numbers and measurement noises.

A machine learning (ML) approach is developed to predict the effect of blend repairs on airfoil frequency, modal assurance criterion (MAC), and modal displacement vectors. The method is demonstrated on a transonic research rig compressor rotor airfoil. A parametric definition of blend geometry is developed and shown to be capable of encompassing a large range of blend geometry. This blend repair geometry is used to modify the airfoil surface definition and a mesh morphing process transforms the nominal finite element model (FEM) to the repaired configuration. A multi-level full factorial sampling of the blend repair design space provides training data to a Guassian stochastic process (GSP) regressor. The frequency and MAC results create a vector of training data for GSP calibration, but the airfoil mode shapes require further mathematical manipulation to avoid creating GSP models for each nodal displacement. This paper develops a method to significantly reduce blended airfoil mode shape emulation cost by transforming the mode shape training data into a reduced basis space using principal component analysis (PCA). The coefficients of this reduced basis are used to train a GSP that can then predict the values for new blended airfoils. The emulated coefficients are used with the reduced basis vectors in a reconstruction of blended airfoil mode shape. Validation data is computed at a full-factorial design that maximizes the distance from training points. It is found that large variations in modal properties from large blend repairs can be accurately emulated with a reasonable number of training points. The reduced basis approach of mode shape variation is shown to more accurately predict MAC variation when compared to direct MAC emulation. The added benefit of having the full modal displacement field also allows determination of other influences such as tip-timing limits and modal force values.

The Paper presents a novel computationally efficient physics based framework for continuous assessment of cyclic and time dependent damage consumption for Siemens Aeroderivative power turbine components based on actual engine operation. The framework discussed in paper provides the capability for Siemens’ customers to move away from fixed overhaul schedule to a customized schedule which is based on a given gas turbines actual operation and inspection findings. This customized overhaul schedule enables the customers a flexibility to maximize the unit availability and minimize operating costs. Semi-empirical framework discussed in this paper, utilizes dynamic systems theory-based approach to estimate the cyclic & creep damage as a response to transient engine operation; characterized by relevant installed engine instrumentation data from the Engine Health Monitoring system. To estimate damage response through any given complex transient operating cycle, algorithm solves a set of ordinary differential equations (ODEs), that have been calibrated to the engine control and safety instrumentation parameters such as shaft speed, turbine temperatures, pressures etc. by pre-analyzed operating envelope cases. The framework can be setup for predicting accumulated cyclic and creep damage for all type of turbine components (Aerofoils, disks, casings, diffusers etc.), transient stress state complexity (in-phase, out-of-phase, uniaxial, multiaxial stress profiles) and is capable to handle unit specific ramp rates, start-up times, restart/cooldown effects specific and random changes in load, history. The framework for discussion in this paper has been demonstrated as applied to the stage-1 blade of an A-35 RT62 power turbine as an example.

The Reynolds number effects on performance of a Controlled Diffusion Airfoil (CDA) geometry in a linear compressor cascade setup are presented. A set of highly-resolved Large Eddy Simulations at Reynolds numbers from 300,000 to 1,000,000 at a Mach number of 0.7 are performed, with one of the simulations including incident wakes at engine relevant reduced frequency of 1.56. The focus of this work is to investigate loss variation of a modern blade geometry by applying both the novel entropy analysis introduced by (Zhao & Sandberg, GT2019-90126) and the mechanical work potential analysis by Miller (GT2013-95488) in the form presented by Leggett et al. (GT2019-91253). The loss analysis, applied both to the whole blade, and to carefully selected subvolumes, including suction and pressure surface boundary layers, highlights regions of increased loss production and the effects of Reynolds number on the location and strength of different flow processes, such as transition and separation. The results are further broken down to show the contribution to loss of individual terms of the analyses and the changes that occur due to changes in Reynolds number. The effects of incident wakes are also captured using this analysis technique and detailed in the paper.

Rim-driven hub-less fans have newly emerged as the most compact type of axial flow fans, which permits flexible configuration arrangements, large relative flow area and low-noise level operation. However, previous publications on rim-driven axial flow fans are rarely found in the open literature, and the flow mechanism and design principle of such promising fans haven’t yet been well-understood and established. This paper has been focused on a preliminary study of the rim-driven axial flow fan design and flow mechanism. A design method of the rim-driven fans is proposed on the basis of the isolated airfoil scheme and the variable circulation rule. It is further incorporated into a FORTRAN code and suited for designing the rim-driven hub-less fans of low-pressure levels. For validation purpose, a conventional hub-type fan is redesigned with the developed method and its flow behavior and overall performance are investigated numerically. A parametric study on the designed fan is further conducted respectively for the tangential velocity difference at mean span, circulation exponent and sweep angle and their influence on the fan flow characteristics and overall performance are explored and highlighted. On such a basis, the developed design method for the rim-driven axial flow fan is further improved. In comparison with the conventionally designed fan at identical rotating speed, significant comprehensive gains are arising from the redesigned fan of hub-less configuration: the overall pressure rise and static pressure efficiency is enhanced respectively by 6.2% and 11.5%, whereas the diameter of the fan is reduced by 12.5% simultaneously. It is demonstrated that the rim-driven hub-less configuration is promising for the enhancing the fan overall performance with even reduced dimensions.

The drive for ever higher aircraft efficiency inevitably leads to a reduction in core compressor size; over the past 30 years the height of the rear stages of a typical aeroengine has reduced by 40%. This trend will continue; a further reduction of almost 50% is expected by 2050. It is shown in this paper that the two greatest sources of increased profile loss as core size is reduced comes from the mechanical and manufacturing constraints placed on the maximum and trailing edge thickness. Furthermore, mitigation strategies are presented which successfully minimized or eliminated these loss sources. An experimental study showed that local modification to the trailing edge shape from a semi-circle to a more elliptical shape almost entirely eliminated the increase in trailing edge loss associated with a reduction in compressor size. Elliptical trailing edges improved loss by both narrowing the wake and increasing the base pressure. Despite the local nature of the modification, the elliptical trailing edges had a powerful effect on flow turning. This was accounted for by redesigning each design to achieve the specified exit flow angle. A set of “as manufactured” small core blades was created by combining measured data from existing manufacture methods with airfoils scaled to ensure minimum thicknesses necessary for mechanical integrity along the blade were not breached. Core sizes down to 40% of current designs were run in CFD. It is often assumed that the loss in efficiency as core size is reduced is unavoidable. However, it was shown that approximately half of this is directly due to the increase in thickness to satisfy tolerancing. Part to part deviation has negligible effect at all but the smallest compressor size tested (at 40% of current values).

Effects of surface roughness on the transition of flat-plate boundary layers under a high-lift airfoil pressure gradient with low incoming freestream turbulence level have been investigated. Time-resolved streamwise and wall-normal velocity fields with surface roughness values of R a /C = 0.065·10 −5 , 4.417·10 −5 and 7.428·10 −5 have been measured at a fixed Reynolds number of 5.2·10 5 and freestream turbulence intensity of 0.2%. For the reference Smooth surface of R a /C = 0.065·10 −5 , a laminar separation bubble forms from about 64% to 83% of the chord length. Displacement thickness increases downstream of separation and then decreases during the transition (reattachment), and momentum thickness increases due to the vortices shed from the separation bubble. Increasing surface roughness has little impact on the laminar boundary layer separation onset but reduces the height and length of the separation bubble and induces earlier transition. For R a /C = 4.417·10 −5 , displacement thickness during transition is slightly thinner and the overall momentum deficit is slightly lower than those for R a /C = 0.065·10 −5 . For R a /C = 7.428·10 −5 , the separation bubble becomes hardly visible as the transition mode approaches the attached mode, and turbulent mixing by the wall-bounded turbulence becomes dominant. In addition, the portion of turbulent wetted area increases due to earlier transition, and momentum deficit increases more rapidly in the turbulent wetted area. Thus, the overall momentum deficit for R a /C = 7.428·10 −5 is larger than that for R a /C = 0.065·10 −5 .

Design methodologies for axial compressor airfoils have undergone significant changes over the past decades. While inverse design methods have played a significant historical role, today they are mostly replaced by direct methods. Inverse methods do impose either the desired pressure or velocity distribution and search for the corresponding blade profile, in contrast to direct methods which modify directly the blade shape to reduce losses. Inverse methods therefore require the designer to know pressure or velocity profiles which provide low losses, and are as such mostly effective only in the hands of an experienced designer. Inverse methods, however, pose some advantages: through setting velocity profiles which feature good off-design performance, the computational cost for the design of profiles can be significantly reduced compared to direct methods, which require to simulate multiple operating points. Additionally, inverse methods offer a way to adapt blades for experimental testing if the wind tunnel imposes restrictions on e.g. Mach number, allowing for similar boundary layer conditions. Finally, inverse methods can be used to deduce the blade geometry from measured or published velocity distributions. Within this article, we aim to verify the use of inverse methods by applying more recent optimisation techniques to the inverse problem. Specifically, we test the performance of an inverse method that uses a gradient based technique to solve the inverse problem. The merits of the inverse method are investigated for different use cases. It is found that conventional, direct design methods are preferred for design improvement, although more expensive. The inverse method is, however, well-suited for adapting existing profiles to altered operating conditions, and for reproducing the blade shape based on published data.

Engine operating cost contributes a major share to an aircraft’s direct operating cost. Thus, the knowledge of the current and future state of their engines is a major concern to any airline operator. To be able to schedule shop visits, state-of-the-art diagnostic and prognostic tools including CFD methods are employed. These RANS-based turbulence and transition models are used to predict the overall efficiency and operational behavior of the engine components. Aerofoil surfaces undergo dynamic change during operation and surface roughness increases in complex non-homogeneous ways due to corrosion, erosion, and fouling processes; depending on the engine component and the environmental condition encountered. The influence of real fouling based roughness on the boundary layer transition is investigated experimentally and numerically within this study. For this purpose, the rotor midspan from the second HPC rotor of the CFM56 is used as the basis for experimental and numerical investigations. Realistic fouling based roughness is applied and investigated both in a cascade tunnel and a low speed compressor rig. The results shown here indicate that laminar boundary layers and their transition to turbulence must be included in the RANS model combination used. Furthermore, it is necessary to consider roughness effects in the respective turbulence and transition model. While the consideration of roughness for the turbulence models has already found wide acceptance, the results in this work motivate the additional extension of the transition model to include roughness effects.

Turbine Vane Frames (TVF) are a way to realize more compact jet engine designs. Located between the high pressure turbine (HPT) and the low pressure turbine (LPT), they fulfill structural and aerodynamic tasks. When used as an integrated concept with splitters located between the structural load-bearing vanes, the TVF configuration contains more than one type of airfoil with sometimes pronouncedly different properties. This system of multidisciplinary demands and mixed blading poses an interesting opportunity for optimization. Within the scope of the present work, a full geometric parameterization of a TVF with splitters is presented. The parameterization is chosen as to minimize the number of parameters required to automatically and flexibly represent all blade types involved in a TVF row in all three dimensions. Typical blade design parameters are linked to the fourth order Bézier-curve controlled camber line-thickness parameterization. Based on conventional design rules, a procedure is presented, which sets the parameters within their permissible ranges according to the imposed constraints, using a proprietary developed code. The presented workflow relies on subsequent three dimensional geometry generation by transfer of the proposed parameter set to a commercially available CAD package. The interdependencies of parameters are discussed and their respective significance for the adjustment process is detailed. Furthermore, the capability of the chosen parameterization and adjustment process to rebuild an exemplary reference TVF geometry is demonstrated. The results are verified by comparing not only geometrical profile data, but also validated CFD simulation results between the rebuilt and original geometries. Measures taken to ensure the robustness of the method are highlighted and evaluated by exploring extremes in the permissible design space. Finally, the embedding of the proposed method within the framework of an automated, gradient free numerical optimization is discussed. Herein, implications of the proposed method on response surface modeling in combination with the optimization method are highlighted. The method promises to be an option for improvement of optimization efficiency in gradient free optimization of interdependent blade geometries, by a-priori excluding unsuitable blade combinations, yet keeping restrictions to the design space as limited as possible.

For the past seven decades, a compressor aerodynamicist has developed various methodologies to design, analyze, and simulate compressor stages. In compressor design, three major subsequent steps can be identified: the one-dimensional mean-line methodology, the two-dimensional through-flow analysis, and the three dimensional computational fluid dynamics. One of the interconnecting threads, between these various x-dimensional analysis, is the compressor blade profile shape. This shape, of known and controllable geometric parameters, is usually accompanied by, or related to, loss models and known flow physics, either defined by theory or through experimental test. In this paper, a novel mathematical approach is described to define axial compressor airfoil profile shapes. These shapes, developed in a Cartesian coordinate system, can be used to create Double Circular Arc, Multiple Circular Arc, and a hybrid combination of the two types. The proposed methodology, based on the mathematics of circles, can be easily applied using generalized software such as Python or MATLAB, or be embedded in specialized engineering design software. In doing so, researchers and engineers can create compressor airfoil shapes which are consistent and flexible with respect to geometric parameter manipulation. Full details of the formulas, with respect to the camber line definition and the calculation of the profile intrados and extrados, are presented. A URL link to an equivalent MATLAB code, and a specialized engineering software, has been provided for those researchers that wish to apply the formulations and review its use.

The scope of this paper is to provide clarity over fundamental effects of non-axisymmetric endwall contouring in a highly-loaded compressor tandem stator. Specifically, the focus of the research will be the influence that aerofoil optimization and end-wall contouring have on each other, and how a combined optimization of the two simultaneously affects the final design of the optimized geometry. The reference geometry used in this research is a standard tandem vane arrangement with smooth axi-symmetric endwalls and designed to represent a datum stage configuration for future investigations of such blade geometries, experimental work on a low-speed research compressor being a next step. The optimization was performed using an in-house developed routine coupled with Auto-Opti, an automated optimizer based on an evolutionary strategy algorithm, developed by the DLR Institute of Propulsion Technology. The entire optimization was conducted solely on the stator, modeled as an annular cascade. This paper reports about a thorough flow field analysis of the optimized geometry in order to understand local mechanisms occurring with non-axisymmetric contouring in the tandem stator passage flow field and its overall performance. Furthermore, it describes and explains the effect that the aerofoil optimization has on the contoured hub surface shape, compared to an optimization process, which is only applied to the hub contouring. In particular, the results clarify how endwall flow field improvements reduce the degree of required aerofoil deformation, and show that the new blade shape is better adapted to the contoured endwall. The 3D endwall flow field has been investigated in order to understand how the optimized geometry modifies the magnitude of the cross-passage flow and reduces the size of the trailing edge corner vortex of the front and rear tandem vane. The paper concludes with some guidelines on how endwall contouring and near endwall aerofoil section design and optimization can be applied most effectively.

The StreamVane™ swirl distortion generator, developed by Virginia Tech, can efficiently reproduce the boundary layer of an airframe or duct found in boundary layer ingesting (BLI) aircraft. Due to manufacturing limitations, the vanes within StreamVanes induce unsteady, vortical wakes, commonly known as a von Karman vortex street. This paper investigates the use of a commercial URANS code and SST turbulence model to predict the vortex shedding frequency from the vanes. The objective was accomplished in two main tasks. First, the CFD methodology was validated by modeling the fluid dynamics of a linear cascade experiment done by the von Karman Institute. Second, the same methodology was applied to airfoils used in StreamVane design to calculate the shedding frequency as a function of turning angle and TE thickness. It was predicted that an increase in turning angle exponentially increased the shedding frequency while an increase in TE thickness exponentially decreased the shedding frequency. The results provided a correlation between the shedding frequency and airfoil characteristics in StreamVanes as well as various turbomachinery components.

This paper applies Co-flow Jet (CFJ) active flow control airfoil to a NREL horizontal axis wind turbine for power output improvement. CFJ is a zero-net-mass-flux active flow control method that dramatically increases airfoil lift coefficient and suppresses flow separation at a low energy expenditure. The 3D Reynolds Averaged Navier-Stokes (RANS) equations with one-equation Spalart-Allmaras (SA) turbulence model are solved to simulate the 3D flows of the wind turbines. The baseline wind turbine is the NREL 10.06m diameter phase VI wind turbine and is modified to a CFJ blade by implementing CFJ along the span. The baseline wind turbine performance is validated with the experiment at three wind speeds, 7m/s, 15m/s, and 25m/s. The predicted blade surface pressure distributions and power output agree well with the experimental measurements. The study indicates that the CFJ can enhance the power output at the condition where angle of attack is increased to the level that conventional wind turbine is stalled. At the speed of 7m/s that the NREL turbine is designed to achieve the optimum efficiency at the pitch angle of 3°, the CFJ turbine does not increase the power output. When the pitch angle is reduced by 13° to −10°, the baseline wind turbine is stalled and generates negative power output at 7m/s. But the CFJ wind turbine increases the power output by 12.3% assuming CFJ fan efficiency of 80% at the same wind speed. This is an effective method to extract more power from the wind at all speeds. It is particularly useful at low speeds to decrease cut-in speed and increase power output without exceeding the structure limit. At the freestream velocity of 15m/s and the CFJ momentum coefficient C μ of 0.23, the net power output is increased by 207.7% assuming the CFJ fan efficiency of 80%, compared to the baseline wind turbine due to the removal of flow separation. The CFJ wind turbine appears to open a door to a new area of wind turbine efficiency improvement and adaptive control for optimal loading.

U.S. Navy marine gas turbine engines serve as primarye and auxiliary power sources for several current classes of ships. Early observations noted in the 1960s and 1970s revealed severe corrosion attack on the first stage blade and vane components of a shipboard marine gas turbine engine that caused engine failure after only several hundred hours. In gas turbine development, there is always a drive and need to enhance the performance and life of engines. The virtues of using Ni-base superalloys in hot-section components has been well recognized and practiced as a means of substantial increase in turbine-inlet temperature, resulting in improvements in thermal efficiency, durability, and performance of engines. The USN shipboard environment (the marine environment) is high in salt laden air and water, coupled with air and fuel sulfur species that cause aggressive corrosion in gas turbine hot sections. Materials that can function in this environment are considered to be “Marinized”. Higher engine power density and pressure ratios for new engine designs will increase maximum blade, vane, and rotor metal temperatures from a mainly Low Temperature Hot Corrosion (LTHC) regime into both the High Temperature Hot Corrosion (HTHC) and Oxidation Corrosion regions. It is expected that future increased surface combatant loads and operational changes will require increased gas turbine operating temperatures and change the associated operating environment to one where Type I and Type II hot corrosion AND oxidation will be prevalent in newly anticipated operational profiles. The advanced gas turbine upgrade package will include better corrosion and oxidation resistant capability and/or higher temperature capable materials and their associated component overhaul methodologies. New materials need to be created and developed for use in more aggressive environments and higher temperature operations. The main cause of the shorter time between overhauls is the materials deterioration of the engine components associated with the hot section of the engine, e.g. turbine airfoils. The deterioration mechanisms are hot corrosion, with Type 1 hot corrosion mechanism becoming operative at the higher temperatures. The goal of this paper is to evaluate methods to enable running the engine at high power while getting back to the longer mean time between overhauls. The method to achieve the longer time is to evaluate and propose for implementation materials, which can withstand the higher temperatures and at the same time mitigate the operative corrosion mechanisms associated with marine environments.

This paper reports the investigation of a flow control strategy for an S-duct diffusers. The method incorporates stream-wise tubercles, and aims to enhance the performance of S-duct inlets by reducing the size and intensity of separated flow. These devices, bioinspired from humpback whale flippers’ leading edge protuberances, have been shown to be effective in increasing post-stall coefficients of lift of airfoils. In S-duct diffusers, the presence of convex curvature next to the separated region provides an ideal location for the installation of a tubercle-like device. The flow control effectiveness was evaluated by test-rig measurements and computational fluid dynamics (CFD) simulations of the flow in an S-duct at high subsonic flow conditions (Ma = 0.80). The S-ducts were rapid prototyped in plastic using 3D printing. Static surface pressure along the length and total pressure at the exit revealed pressure recovery, total pressure loss, swirl, and the nature of flow distortion at the S-duct exit. CFD simulations used ANSYS FLUENT with a RANS solver closed with the RKE turbulence model. The CFD simulation compared well with the test-rig data and provided useful information on flow mechanism and for understanding flow features. The performance of the baseline and variant with the flow control device was compared and flow control strategy was evaluated.

The performance of a compressor stator airfoil with end-wall injection was studied experimentally and computationally. The geometry was a high-speed, subsonic, linear cascade. The independent variables studied were airfoil incidence angle and mass flow rate of end-wall injection upstream of the stator. The end-wall injection was intended to simulate upstream “leakage” through hardware gaps in the end-walls of gas-turbine engines. The exit of the cascade was interrogated experimentally by a five-hole-probe and a total pressure Kiel probe to provide total pressure measurements, which were used to calculate total pressure loss coefficients at the exit of the test section. Computational studies were completed to examine the end-wall flow physics and entropy generating mechanisms through the stator section. The experimental results showed a distinct decrease in the downstream total pressure field with end-wall injection flow, and the impact of the upstream injection on the stator loss coefficient was not a function of the incidence angle. The computational investigation found that the majority of the end-wall injection’s effect on the downstream total pressure field was observed as an increase in the size of the secondary flows on the suction-side of the stator near the upper end-wall.