bluuebunny/crossref_metadata_2025 huggingface dataset:
FELTMETAL™ Abradable Turbine Seal Materials: Structure and Property Responses to Blade Rub and Oxidation
Authors: Motyka Elaine , Schricker Robert , Ceiler Kelly
Published: October 2022

Abradable seals are a common technology employed to improve turbine engine performance by reducing the gap between blades and the casing in compressor and turbine stages. Because operating conditions and material requirements vary significantly as a function of turbine stage and turbine design, many abradable compositions have been developed. Generally, all abradable seal materials employ voids or pores in the microstructure as a designed-defect to mitigate thermal and kinetic energy from blade rub. Some also have added non-metallic phases to provide other means of maintaining a seal that is sacrificial to the blade tip.

This paper describes the microstructures of specific fiber metal abradable materials known as FELTMETAL™, as well as the distinctly different abradable microstructures of honeycomb material and thermal spray coating. The effect of the porosity and metallic surface area in each type of microstructure on seal abradability and seal oxidation is discussed. In addition, the effect of oxidation by aging at elevated temperatures on strength, ductility, and abradability for several FELTMETAL™ materials is also discussed.

It is proposed that the pore structure aids in mitigating blade contact energy primarily by internal deformation of the metal matrix surfaces and that abradability is improved by formation of protective metal oxide films that reduce ductility without significant loss of strength.
Numerical Investigation of the Effect of Trailing Edge Thickness of Simulated CMC Blades on Loss Profiles
Authors: Miki Kenji , Ameri Ali
Published: October 2022

Ceramic Matrix Composite (CMC) is an enabling material allowing higher turbine inlet temperatures and possibly resulting in better thermal efficiency of gas turbine engines. This is because CMC layers with environmental barrier coating (EBC) have significantly higher thermal limits (∼1755K) compared with the more conventional alloyed metallic blades. CMCs possess a complex fabrication process resulting in different geometrical characteristics than metallic blades (e.g., larger trailing edge thickness, large leading-edge curvature, etc.). It is therefore desirable to assess the aerodynamic performance of the CMC blades using experimental and numerical simulations. To investigate, three different CMC blades with varying trailing edge thicknesses were numerically simulated. Both Large Eddy simulation (LES) with the dynamic subgrid closure as well a recently implemented Reynolds-averaged Navier-Stokes (RANS) coupled with an intermittency function-based transition model were utilized. Two sets of experimental test points with different Reynolds numbers at a given high-free stream turbulence range (Tu = ∼10–13%) were considered. The LES was previously validated by comparing with experimental data. The RANS transition model was validated against the experimental data of the Stripf’s for a turbine blade. The predicted pressure loading profiles, total loss distributions, and integrated loss of the three different CMC blades were computed and comparisons against the experimental data are presented herein. The mixing and boundary layer loss was evaluated by both the RANS and LES and was found that the mixing loss forms a larger proportion of the overall loss as the trailing edge thickness increases.
Shape Optimization of the Laidback Fan-Shaped Film Cooling Hole on Pressure Surface of Turbine Guide Vane
Authors: Jiang Yan , Tao Zhi , Li Haiwang , Zhou Zhiyu
Published: October 2022

The present study employed computational fluid dynamics (CFD) to explore the effect of the shape parameters of fan-shaped holes on film cooling effectiveness, and the optimum values and variation laws of each shape parameter to maximize the averaged film cooling effectiveness were presented. Meanwhile, experimental verification was used in the optimum holes and reference cylindrical holes.

Among the shape parameters of a laidback fan-shaped film hole, the injection angle, the hole diameter and the hole length were fixed as 30 degrees, 0.5 millimeter and 2 millimeters, respectively, and the three shape parameters, the forward expansion angle, the lateral expansion angle, and the shaped-length ratio, were selected as design variables by the Box-Behnken Design (BBD) method. And RSM (Response Surface Methodology) was used to obtain the optimized values of each parameter. Results show that the film cooling effectiveness of the optimized hole (15-20-0.25) was higher than that of the reference cylindrical holes by 336%. The factors influence analysis showed the film cooling effectiveness increased as the forward expansion angle increased, and increased and then decreased with the lateral expansion angle and the shaped-length ratio. The points of the change in trends were when the lateral expansion angle was almost equal to 15 degrees and the shaped-length ratio was nearly 0.475, respectively. Moreover, there was the oscillation of the film cooling effectiveness distribution, possibly due to the asymmetric and irregular distribution of velocity w of the injected coolant along the downstream direction, when the holes outlet area was excessive. In general, the variation laws were not monotonous; therefore, all shape parameters should be considered simultaneously to achieve shape optimization of the laidback fan-shaped film cooling hole on pressure surface of turbine guide vane.
Calculating Cooled Turbine Efficiency With Weighted Cooling Flow Distributions
Authors: Berdanier Reid A.
Published: October 2022

The ability to accurately characterize the thermal efficiency of turbine stages with cooled components is essential to advance turbine technologies. There are many methods for calculating aerodynamic efficiency of turbomachinery components, with most encompassing either single point or simple integrated values at the inlet and exit locations. For cooled turbine stages, there is an added complexity of how to incorporate cooling flow streams. Although these techniques may include multiple cooling flows, a fully-mixed assumption is often invoked without directly accounting for distributions of cooling flow in the annulus. However, the distribution of cooling flows in turbine stages may vary significantly depending on the location and nature of injection. This study addresses how calculated efficiency may vary depending upon how cooling flows are injected and the manner in which these cooling flows are accounted. The analyses utilize experimentally measured cooling flow concentration profiles to motivate a new approach for calculating cooled turbine efficiency with weighted cooling flow distributions. Validation data for a one-stage high-pressure turbine geometry from the Energy Efficient Engine (E3) program are assessed to determine the range of influence for cooling flow distributions. Results show that cooling flows injected at similar axial positions can have unequal influences on integrated stage efficiency based on known distributions of cooling flow and other properties at the turbine stage exit plane. In some cases, deviations approaching half a point of efficiency can be observed relative to traditional equations, and the magnitude of influence is shown to increase as stage total pressure ratio decreases.
The Inter-Bristle Pressure Field in a Large-Scale Brush Seal
Authors: Bowen Joshua P. , Jenkins Matthew R. , Bowsher Aaron A. , Crudgington Peter F. , Sangan Carl M. , Scobie James A.
Published: October 2022

Brush seals promise improvements to the widely used labyrinth seal in regulating turbomachinery leakages. Enhanced resistance to the flow is provided by a static ring of densely packed fine wire bristles that are angled in the direction of rotation and flex to accommodate rotor excursions. A large-scale brush seal was constructed to study the leakage characteristics in direct relation to the pressure field within and surrounding the bristle pack for multiple clearance conditions, therefore developing the understanding of brush seal fluid dynamic behaviour.

The governing parameter controlling leakage behaviour transitioned from pressure ratio for a large clearance, to pressure load for a line-on-line configuration. In all cases, leakage flow converged to an asymptotic value once maximum levels of bristle blow-down and pack compaction were attained. For both clearance configurations, this occurred at a pressure ratio corresponding to that at which axial distributions of pressure converged; equivalent behaviour was noted for the line-on-line configuration with pressure drop. Comparatively small changes were experienced in leakage behaviour and to the inter-bristle pressure field with increasing pressure drop for the line-online brush seal. This indicated that brush seal performance is more influenced by changes in bristle blow-down than bristle pack compaction.
Creep Analysis and Microstructural Evaluation of a Novel Additively Manufactured Nickel-Base Superalloy (ABD®-900AM)
Authors: Bridges Alex , Shingledecker John , Clark John , Crudden David
Published: October 2022

Nickel-base superalloys containing 30 to 50% gamma prime (γ′) volume fraction are typically used in hot section components (e.g. guide vanes or blades) for power generating gas turbines, and suitable time dependent properties are required for long-term elevated temperature operation. Additive manufacturing (AM) has recently been used to develop complex hot-section parts utilizing innovative designs with enhanced cooling features which improve efficiencies by reducing cooling air consumption.

To further explore the opportunity to improve time-dependent AM superalloys, this paper focuses on a fundamental creep study and characterization of a novel nickel-base superalloy (ABD®-900AM) that was manufactured using a laser-based powder bed fusion AM process. The material was subjected to a sub-solvus solution anneal and multi-step aging heat-treatment to produce a bi-modal distribution with ∼35% volume fraction of gamma prime without post-processing hot isostatic pressing (HIP). Microstructural characterization was carried out for the as-built and fully heat-treated structures, and a creep-rupture test program was conducted to study the resultant creep properties. Activation energies and stress exponents in addition to rupture strength and deformation resistance, were compared to traditionally cast IN939 and IN738 materials. After testing, specimens were evaluated using a variety of microscopy tools to determine location and features associated with creep damage.

The optimized chemistry for ABD®-900AM was printed crack free and fully dense in contrast to studies on similar alloys where significant process development and post-build heat-treatments were required. High-temperature mechanical properties in the heat-treated material showed some decrease in creep strength when compared to traditional casting. This strength and rupture life debit was dependent on build orientation, but a considerable increase in creep ductility was observed due to differences in the microstructure when compared with similar AM alloys. Analysis of creep data showed differences in creep mechanisms compared to traditional cast alloys. The relationship between microstructure and creep mechanisms is discussed, and ongoing work to further improve rupture strength through heat-treatment optimization will be highlighted.
Influence of Pressure Gradient on Flame-Vortex Interaction and Flame Stability
Authors: Yalcinkaya Yagiz , Bozkurt Ogeday E. , Gungor Ayse G.
Published: October 2022

This study presents numerical investigations of turbulent premixed bluff-body stabilized flame by emphasizing the influence of pressure gradient on flame-vortex interaction and flame stability for lean combustion applications. Large eddy simulations of four different geometrical configurations, diffuser 3°, diffuser 1.5°, nominal, and nozzle that resulted in mild to strong pressure gradients are presented. Numerical investigations allowed determining the effects of geometry-induced pressure gradient on the flame structure, development of the flame-front vorticity and turbulent structures and flame stabilization. It is shown that the pressure gradient plays a key role for the spatial and temporal development of the flame front vorticity and baroclinic torque. The flow deceleration in diffuser geometries suppresses the flame-induced vorticity mechanisms, which in turn lead to large wrinkle forms of the flame and may lead to local extinctions along the flame front. The favorable pressure gradient in the nozzle geometry, on the contrary, increases the baroclinic torque that restrains the development of the shear layer vorticity and hence prevents local extinctions.
Profile Loss Reduction of High Lift Turbine Blades With Rough and Ribbed Surfaces
Authors: Ananth S. M. , Nardini Massimiliano , Vaid Aditya , Sandberg Richard D. , Vadlamani Nagabhushana Rao
Published: October 2022

Low pressure turbines (LPT) typically operate at low Reynolds numbers, of O(105), resulting in transitional boundary layers on the suction surface. These boundary layers are prone to separation under the strong local adverse pressure gradients on the aft portion of the blade. Intermittent free-stream turbulence, periodic wakes shed by the upstream blades and surface roughness due to gradual degradation of the blades have been shown to suppress the separation bubble on the suction surface of the LPT blade. Although this generally leads to a profile loss reduction, some of the benefit is offset by a loss increase associated with a larger turbulent wetted area. In this work, we explore a strategy where the losses in both the transitional and turbulent boundary layers can be reduced. In particular, we employ surface roughness in the transitional regime to trip the boundary layer and reduce the separation bubble related losses. Subsequently, we explore the possibility of utilizing riblets in the turbulent regime to further reduce the losses due to the turbulent wetted area. To investigate the efficacy of this ‘rough-riblet blade surface’, high fidelity eddy-resolving simulations are carried out on the configuration of a flat surface subjected to streamwise varying pressure gradients. A contoured upper wall is used to mimic the pressure distribution typically encountered on the suction surface of a high lift turbine blade. The Boundary Data Immersion Method is used to represent different riblet shapes and roughness elements, while a digital filtering approach is used to impose free-stream turbulence at the inlet. The influence of different riblet geometries on skin friction drag of transitional and turbulent boundary layers under streamwise varying pressure gradients are discussed in detail. The effect of different riblet shapes in conjunction with the roughness elements on the passage losses is shown through boundary layer integral parameters and Reynolds stresses.
An Experimental Test Case for Transonic Low-Pressure Turbines - Part 2: Cascade Aerodynamics at On- and Off-Design Reynolds and Mach Numbers
Authors: Lopes G. , Simonassi L. , Torre A. F. M. , Patinios M. , Lavagnoli S.
Published: October 2022

In this two-part paper, a novel test-case for transonic low-pressure turbines (LPT) is presented. The current study is a comprehensive report on the design, commissioning and testing of a high-speed LPT cascade. Part II reports the characterization of the aerodynamics at on- and off-design flow conditions.

A detailed analysis of the steady aerodynamics of the highspeed low-pressure turbine blade is presented for a range of engine representative outlet Mach numbers from 0.70 to 0.95 and Reynolds numbers from 65,000 to 120,000. The blade 3D aerodynamics are characterized using an innovative set of traversable blades, enabling high-resolution radial measurements. A novel method to estimate the location of separation-reattachment based on pneumatic tap measurements is presented. The separation on the blade suction side is strongly influenced by the Reynolds number at the lower Mach numbers, while open separations were observed at transonic exit conditions independently on the Reynolds number.

Downstream measurements by means of a five-hole probe, and pressure taps located in the passage endwall are employed to study the secondary flow development and structures at the cascade outlet. The results show that the losses follow different trends at high and low Reynolds numbers. The profile losses at Re = 70k decrease with increasing Mach number, contrary to what is observed for Re = 120k. The minimum secondary flow losses are found for an off-design condition with lower Mach number with respect to the nominal. The off-design comparisons presented in this paper indicate that at low Reynolds, operating at transonic outlet Mach numbers leads to beneficial effects on the performance.
Potential and Evolution of Miniatures Compressed Air Energy Storage Plants Based on Impulse Turbine
Authors: Al Jubori Ayad M. , Al-Sadawi Laith A. , Biedermann Till , Alfarawi Suliman
Published: October 2022

This paper describes the work carried out to develop an impulse turbine for miniatures compressed air system. This study hypothesizes the question; what is the effect of combining an impulse turbine loss model into a compressed air energy storage system analysis? The miniatures power system has lower mass flow rates which lead to a small turbine size. The miniature impulse turbine has relatively low efficiency and is highly sensitive to operating conditions at a low mass flow rate due to all losses in terms of passage, trailing edge, incidence, and clearance becoming higher amounts compared to the total losses of the percentage foundation.

The development of a novel impulse turbine configuration is presented based on one-dimensional design and three-dimensional simulations. The impulse turbine in single-stage configuration was designed and analyzed for a range of operating conditions in terms of pressures, temperatures, mass flow rate, and rotational speeds. The simulations results showed that the maximum efficiency and power were 65.93% and 4.019 kW respectively with a mass flow rate of 0.2 kg/s. The energy analysis revealed that the system efficiency was 10.3%. The miniature compressed air energy storage system driven by an impulse turbine can be used to generate electricity for small power applications.
High Pressure Testing With Optical Diagnostics of a Hydrogen Retrofit Solution to Eliminate Carbon Emissions
Authors: Koomen Joris , Dammers Tim , Demougeot Nicolas , Stuttaford Peter , Heinze Johannes , Stockhausen Guido , Fleing Christian
Published: October 2022

Ensuring grid stability is becoming an increasingly large challenge as the amounts of power from new highly fluctuating sources like solar and wind are increasing. Fuel flexible gas turbines which can run on carbon free or low carbon fuels offer grid stability both on the basis of their inertia and the possibility to quickly ramp-up and down.

A consortium led by Thomassen Energy is at work at realizing a cost-effective, ultra-low emissions (sub 9ppm NOx and CO) combustion system for gas turbines in the 1–300 MW output range. A key requirement is fuel flexibility and stable operation from 100% natural gas to 100% hydrogen and any mixture thereof.

This paper describes the results of the second step in the testing cycle of the developed combustion technology. In this step the FlameSheet™ is validated at machine operating conditions. The tests are completed using a state-of-the-art test setup combined with novel optical analysis techniques.

The results show that one of the developed hardware variants can run on a wide range of fuel compositions up to 100% hydrogen without flashback while emissions are below 9 ppm at 15% O2 dry conditions.
Computational and Experimental Study of Film-Cooling Effectiveness With and Without Downstream Vortex Generators
Authors: Lee Chien-Shing , Shih Tom I-P. , Straub Douglas , Weber Justin , Robey Edward
Published: October 2022

Downstream vortex generators that involve a pair of rectangular plates arranged in an open V shape placed just downstream of each film-cooling hole was shown to create flow and vortical structures that entrain lifted film-cooling flow back to the surface and increase its lateral spreading on the surface (GT2020-14317). In this study, computations and measurements were performed to examine the flow mechanisms this vortex generator induces to improve film-cooling effectiveness of a flat plate with the cooling flow emanating from one row of inclined holes. Parameters studied include blowing ratio (BR = 0.75, and 1.0) and temperature ratio (TR = 1.07 and 1.9). The computational study is based on steady RANS closed by the shear-stress transport (SST) turbulence model with and without conjugate analysis. The experimental study was conducted by using a conjugate heat transfer test rig with a plenum, where cooling flow is introduced. Measurements made include velocity and temperature profiles upstream and downstream of the film-cooling holes as well as temperature at several locations on the hot and cold sides of the film-cooled-flat plate. The computational study was validated by comparing computed results with those from measurements at BR = 0.75 and 1.0 and TR = 1.9. Computational and experimental results are presented to show the effects of BR and TR on the flow structures and how those structures improve the effectiveness of film cooling with and without the downstream vortex generators and with and without conjugate heat transfer.
Upgrades to the Single Stage Axial Compressor and Fan Facility for Low Pressure Ratio and Boundary Layer Ingesting Fan Research
Authors: Stephens Julia E. , Hughes Chris
Published: October 2022

Next generation aircraft engine efficiency goals will be met through high bypass ratio engines. To achieve these goals, a combination of small core and low pressure ratio fan technologies will be necessary. Additionally, advanced propulsion airframe integration concepts, such as boundary layer propulsion, and distributed electrically driven fans with short inlets or wing integration strategies, are beginning to be investigated for their potential fuel burn savings.

In order to realize the potential fuel savings from these technologies, it is necessary to have testing capabilities for the components. The NASA Glenn Research Center (GRC) Single Stage Axial Fan and Compressor Facility (also called the GRC W-8 Facility) has recently undergone several significant upgrades in order to effectively test low pressure ratio fans and various inlet distortions. This paper will outline various challenges to testing low pressure ratio fans and the methods implemented to ensure that the W-8 facility is capable of performing these tests and acquiring high quality data. The upgrades include significant piping upgrades to reduce pressure loss and turbulence, capability for inlet total pressure and swirl distortion, increased test section length including a longer shaft, a two component balance, an upgraded rotating data system, and a new temperature characterization to reduce uncertainties in thermocouple measurements.
Experimental Study on Heat Transfer in a Rotating, Ribbed Serpentine Cooling Channel With Bleed Holes
Authors: Zhu Chenghua , Wen Jie , Xu Guoqiang , Chen Yanan
Published: October 2022

The effect of rotation on heat transfer in a typical turbine blade middle-chord channel is experimentally investigated. The channel is characterized by ribs, three-pass serpentine passages and bleed holes. Cooling air is radial outward in the first passage and the third passage, and is radial inward in the second passage. The bleed holes were placed on trailing side (pressure side) of the first passage and the second passage, and on both trailing side and leading side (suction side) of the third passage. An additional small stream of cooling air is injected at the second 180-deg hub turn to compensate the insufficient cooling flow rate in the third passage. The major inlet Reynolds number (based on the inlet of the first passage inlet) and rotation number are varied from 17000 to 33000 and 0 to 0.43, respectively. Besides, the mass flow rate ratio (MR, the mass flow rate of second stream to the major coolant) is varied from 0 to 0.2. The wall temperatures were measured by thermocouples embedded behind the copper plates to obtain the regional average heat transfer coefficients. The result shows that compared with the traditional square U duct, the rotation effect on heat transfer is apparently suppressed in current channel with more realistic blade configurations. According to the experimental data, the overall averaged heat transfer in the first and second passage are respectively weakened by 6.2% and 2.4%. Under the influence of the second stream, rotation elevates heat transfer in the third passage, whereas a heat transfer reduction is observed in the second passage with ascending MR, indicating that high MR case (MR = 0.2) is not recommended whether from saving cooling air or improving the general heat transfer of the cooling channel.
Leakage Flow Impact on Shrouded Stator Cavity Flow Topology and Associated High Speed Axial Compressor Stage Performance
Authors: Babin Cedric , Ottavy Xavier , Fontaneto Fabrizio
Published: October 2022

This paper presents the investigation of realistic cavity leakage flow for the von Karman Institute for Fluid Dynamics H25 axial compressor stage, equipped with a shroud cavity of average Reynolds number Rer = 2.2 * 106 and average aspect ratio G = 0.06. On top of overall performance measurements, time averaged data from unprecedented experimental dataset are used to characterize the stator hub shroud cavity flow field and its interaction with the power stream. The sensitivity of the stage performance and stability as well as the cavity flow topology is evaluated against injection conditions, compressor throttling and compressor speed. Parametric steady RANS simulations of the stage under injection are used to support the findings and state on the relevance of such approach in the preliminary design phase of axial compressor components.

As an example, a stage efficiency reduction of 0.97% for a leakage fraction of 0.56% at design point is retrieved. This performance drop is attributed to the effect of cavity flow on: the blockage ratio, the boundary layer skewness, and total temperature increase at stator row inlet. It is also presented that there is a mutual interaction between the cavity geometry and the cavity flow field organization. The injection homogenizes the cavity flow field and the associated pressure gradient. The various operating conditions presented also demonstrate that the cavity flow studied is, on average (i.e. time), sensitive to changes in compressor rotation speed and changes in stage loading.
A GTC-CGC Health Assessment Method for Gas Turbine Engine
Authors: Ao Ran , Cao Yunpeng , Luan Junqi , Han Xiaoyu , Li Shuying , Yan Li
Published: October 2022

Due to wear, fouling, corrosion, and other factors performance degradation of compressor and turbine seriously affects the health state of gas turbine engines. However, the health status of gas turbine engines is inconsistent under the variances of the external environment and operating load, which brings great challenges to health assessment and remaining useful life predictions. To solve this problem, a grey target calculation and cloud gravity center (GTC-CGC) health assessment method is proposed. The three-dimensional data normalization process is to characterize the different initial states between samples. The empirical signal-to-noise ratio and entropy weight are used to calculate the subjective and objective weights of health indicators, and convex optimization is used to realize the fusion assessment. A health assessment model was established by the sliding time window to track and calculate the real-time health state of the gas turbine engine. Finally, health assessment and remaining useful life prediction tests were conducted on turbofan engines using the NASA C-MAPSS dataset. The experimental results show that compared with the self-organizing neural network, the monotonicity, robustness, and trend of the health assessment results were improved by 0.3216, 0.0843, and 0.0355 respectively; Based on the GTC-CGC health assessment results, a third-order linear regression algorithm was used for turbofan engines remaining useful life prediction, the prediction score is 265, and root mean squared error is 26.1542, which is equivalent to the prediction accuracy of mainstream intelligent prediction methods such as convolutional neural network and short-term memory neural network.
Response of Separated Boundary Layers to Steady and Pulsated Flow Injection in Transonic Internal Flows
Authors: Nowak Hunter D. , Lluesma-Rodriguez Federico , Rahbari Iman , Clark John P. , Paniagua Guillermo
Published: October 2022

A new class of power generation devices that experiences increased losses due to bulk flow separation in segments of their expected in-flight regime is emerging. As such, active flow control becomes increasingly relevant to mitigate these losses and reclaim the entire flight envelope. This study explores the effect of flow injection on transonic flows experiencing bulk separation. Reynolds-Averaged Navier-Stokes simulations of a 3D wall-mounted hump at low Reynolds numbers are conducted to assess the response of transonic bulk separation to flow injection. Unsteady simulations are performed to understand the differences between slot and discrete port injection and determine optimum forcing frequencies. Discrete ports require higher pressures to overcome the momentum deficit associated with the smaller injection area relative to the width of the domain. Steady and unsteady injection are found as viable strategies in mitigating the extent (or appearance) of bulk separation. Experiments are conducted with discrete injection for a range of Mach and Reynolds numbers. The response of the bulk separation to said injection is evaluated by analyzing both local pressure measurements and schlieren imaging. The study shows that the required pressure of injection is strongly correlated to the length scale of the uncontrolled separation. With Large Eddy Simulations, the flow separation and frequency content within the separated region can be reasonably predicted. This study aims to take further steps to establish guidelines for applying flow control to the emerging class of power generation devices experiencing losses from bulk separation.
On the Influence of Stator-Rotor Radial Gap Size on the Fluid-Dynamic Performance of Mini-ORC Supersonic Turbines
Authors: Cappiello Alessandro , Majer Matteo , Tuccillo Raffaele , Pini Matteo
Published: October 2022

Radial-Inflow Turbines are considered the most suited expanders for waste heat recovery via high-temperature mini-organic Rankine Cycle (ORC) turbogenerators thanks to high compactness, large expansion ratio handled by a single stage, and comparatively low weight. Reaching high efficiency in these machines is however a formidable challenge, as they are bounded to operate with expansion ratios exceeding 40.

Although scarcely investigated in the published literature, the size of the stator-rotor radial gap is a key design parameter as it has a large influence on fluid-dynamic performance, manufacturing, and mechanical integrity. In addition, the working conditions of the turbine are such that the stator operates with highly supersonic flows in the non-ideal thermodynamic regime, making the flow pattern in stator-rotor radial gap, which can be regarded as an area-decreasing channel, very complex. Under these conditions, the radial gap size could impact the stage efficiency up to few percentage points. The paper presents a study aimed at investigating the impact of variable radial gap on the fluid-dynamic performance of radial-inflow turbines for high-temperature mini-ORC power systems. The reference turbine is a supersonic machine for laboratory experiments under realization at Delft University of Technology, referred to as ORCHID turbine. First, a theoretical analysis is carried out to identify the relevant non-dimensional parameters governing the flow physics in the gap. Then, the effect of the radial gap size on the fluid-dynamic performance of the ORCHID turbine is assessed by means of RANS and uRANS computations. The results show that the change in radial gap size leads to a redistribution of expansion ratio between vaned and vaneless part of stationary component via a substantial change of the stator trailing edge flow structures, which, in turn strongly affects stator loss and stage efficiency, leading to 8% points reduction.
Predicting Separated Flow in a Three-Dimensional Prediffuser for Combustor Applications Using Improved Numerical Techniques and Workflow
Authors: Norman Adam , Arguinzoni Carlo
Published: October 2022

After the last stage of the high-pressure compressor of gas turbine engines, swirling air enters a prediffuser which attempts to straighten out the flow and convert the dynamic pressure to a static pressure. Before the combustion chamber there is an aerodynamic cowl that, typically, separates the air into three parts. The main path leads to the combustor and supplies air to the fuel nozzles, while the secondary paths branch to the inner and outer diameter shroud passages to be used for cooling processes.

Numerical tools are widely used in prediffuser design to find the optimum design that converts a large portion of the dynamic pressure to static pressure, with the target to avoid separation. Typical analysis today for combustor aerodynamics are done with large eddy simulations (LES), but for the prediffuser design most of the analysis are done using a Reynold’s average Navier Stokes (RANS) approach.

The combustor aerodynamic community today, rely on hand-crafted, structured, hexahedral meshes to solve flows inside the prediffuser. Creating a structured mesh can be very time consuming and could delay design evolution depending on the complexity of the prediffuser. Due to the flow complexity at the exit of the prediffuser and the impact that could have on flow split predictions provided to the combustor, Reynold’s stress turbulence models (RSM) are typically used. The combined effort of creating structured meshes, with the difficulty in converging results with RSM can negatively impact design times.

The effort presented here investigates modeling a true three-dimensional diffuser from Cherry et al. using a Poly-Hexcore mesh topology and the Generalized k-Omega (GEKO) turbulence model. The unstructured Poly-Hexcore meshing workflow does not require any manual meshing operations, utilizes parallel meshing, and can be fully automated to simplify the mesh generation process. The GEKO turbulence model has the flexibility to cover a wide range of flows using free coefficients and can be setup to allow robust capturing of prediffuser flow characteristics. Numerical results are compared to experimental data gathered by Cherry et al. and current numerical best practices.
Influence of the Convergent-Divergent Nozzles Arrangement Parameters on Aerodynamic Losses and Overall Supersonic Turbine Performance
Authors: Moroz Leonid , Burlaka Maksym , Yudin Aleksandr
Published: October 2022

Single-stage axial pressure compound and two-row velocity compound supersonic turbines have a very high specific work, which allows for high power levels at high-pressure ratios and small flow rates in a very compact dimensional envelope. The latter is extremely desirable in certain types of Liquid Rocket Engine (LRE) turbopumps. Besides, due to their compactness, mechanical simplicity, and low cost, such turbines are favorable in high-pressure ratio Organic Rankine Cycles (ORC), gas expanders in chemical and technological processes, various mechanical drive applications, etc. However, despite their mechanical simplicity, the supersonic aerodynamic effects in such turbines are very complex. There are a number of design and analysis methodologies for such turbines, which are currently used in the industry. Nevertheless, some important nozzles arrangement parameters are still insufficiently investigated. It is known that there is an influence of the principle of drilled nozzles arrangement in a cascade and (round cross-section) convergent-divergent nozzles overlapping on overall turbine efficiency. Though, the data were either not formalized in some sort of loss model or formalized in a simplified formula that does not take into account meridional nozzle angle and the diameter where the nozzles are arranged (turbine mean diameter). The paper describes the study of the qualitative and quantitative influence of the mentioned above parameters on aerodynamic losses in drilled convergent-divergent nozzles and overall supersonic turbine performance. Supersonic aerodynamic effects and their influence on turbine performance were determined by employing the state-of-the-art commercial CFD code. Respective mesh sensitivity studies and domains architectures selection considerations are presented. The formalization of the studied nozzle parameters is discussed. The methods used for the determination of the parameters dependencies minimizing the number of CFD calculations are explained. The figures with visualizations of the critical results and numerical values are provided. The obtained loss models are described. The obtained loss models were incorporated into a turbine design and analysis program and then compared with experimental data. The utilization of the improved loss model provides the opportunity to design high-efficient supersonic turbines with conical drilled nozzles and satisfy specific design constraints with the use of nozzle arrangement parameters.
An Improved Preliminary Design Tool for Turbomachinery Blades Using Van der Pol Based Reduced-Order Model for Non-Synchronous Vibrations
Authors: Hollenbach Richard , Kielb Robert , Hall Kenneth
Published: October 2022

Starting from an oscillating cylinder, this paper navigates the shift of a Van der Pol based Reduced-Order Model from that simple case to airfoil and turbomachinery blade. The structure will be free to vibrate under the influence of a von Karman vortex street. The first degree of freedom tracks the oscillatory motion of the vortex shedding utilizing a combined Van der Pol and a Duffing equation, previously shown to better match the flow physics. The other degree of freedom is a cylinder or pitching blade with stiffness, mass, and damping, making the coupled system one of fluid-structure interaction. A third degree of freedom is added in a separate configuration to provide the airfoil with both pitching and plunging motion. Using this model to study the time-history of the fluid and the structure oscillation, additional parameters are extracted to understand the underlying mechanisms of frequency lock-in and limit cycle oscillation. The coefficients in the model are tuned to match limit cycle oscillation data captured in experiments. Then, the phase shift between the vortex shedding and the structural motion is calculated when the former locks-on to the latter, and then unlocks. Also, the work done per cycle of vibration is analyzed from the contributions of the mass, spring, and damping forces to determine the dominant contributor when locking-in versus unlocking. This model is a more accurate case for turbomachinery applications compared to the previous models. This model, especially with the third degree of freedom, will serve as the preliminary design tool for engine manufacturers to use for preventing Non-Synchronous Vibrations in turbomachinery.
Design Optimization of Blade Tip in Subsonic and Transonic Turbine Stages - Part II: Flow Physics and Augmented Aerothermal Integral Objective Function
Authors: Duan Penghao , He Li
Published: October 2022

In Part I, a companion paper of the two-part article, two turbine stages working with the same Reynolds number (8.8 × 105), flow coefficient, loading coefficient and reaction, but the different exit Mach number (Mexit = 0.7 and Mexit = 1.1, respectively) are designed and optimized to provide a direct contrast between the high subsonic and the transonic turbine tip. In Part II, further analysis are carried out primarily aimed at addressing the issues of interest arising from Part I:

a) to understand the driving flow physics for the observed contrasting aerodynamic efficiency sensitivities of the two stages with respect to squealer height.

b) to seek a more suitable heat transfer objective function for the tip aero-thermal design optimization, given the seemingly strong conflicts among the heat transfer objective functions.

The advanced post-processing techniques are adopted to identify and quantify vortical structures in the rotor tip region. Two vortices, the Pressure Side Vortex (PSV) and the Casing-driven Cavity Vortex (CCV), are found to have major influence on the flow pattern within the squealer tip.

On the aerodynamic side, the transonic squealer is shown to be somewhat less effective than the subsonic one (Part I). The OTL mass flow rate in the transonic stage is less sensitive to the squealer height. The corresponding reduced blockage for the transonic squealer may be attributed to two factors: 1) the vena-contracta formed by the PSV as it leaves the squealer cavity induces smaller size of separation in the supersonic flow. Thus, the blockage effect and the mass flux deficit are reduced in the transonic stage; and 2) the choking of the OTL at the rear portion of the transonic tip acts as a limiter for the leakage mass flow. On the heat transfer side, the two vortices also have major signatures on the cavity surfaces, mainly by impingement. For the PSV, the impingement impact is mainly on the cavity floor. For the CCV on the other hand, its impact is mainly on the inner side wall of the suction side rim. The overall linear variations of the heat load with the squealer height primarily correspond to the linear increase of the cavity side wall surface. The vortical structure is found to be influential: the higher slope of increments in the transonic stage is due to the stronger and more persistent CCV with a clear signature on the large portion of the suction side cavity wall at a large squealer height.

In addressing issue b), the coolability weighted non-uniformity parameter is proposed to integrate the local heat transfer and the coolability. The proposed coolability augmented objective function is shown to provide more consistent Pareto fronts for heat transfer non-uniformity, than those adopted in Part I. It thus should serve well as an enabler to help practical applications of blade tip aerothermal design optimizations.
A Numerical Investigation of Sweeping Air/Mist Jet Film Cooling Through a Passive Fluidic Oscillator, Part I: Steady vs. Sweeping Air-Only Jets
Authors: Abdelmaksoud Ramy , Wang Ting
Published: October 2022

The sweeping jet can be a promising candidate in film cooling applications since it has a large lateral jet spreading which can be considered an advantage when compared to the regular steady jet film cooling. Fluidic oscillators can generate a sweeping jet without the need for any moving parts. In addition, they can be more conveniently manufactured by additive manufacturing techniques. This two-part paper presents a numerical study to investigate the application of using air/mist sweeping jets in film cooling for protecting turbine airfoils. Part I focuses on validating the computational mode by comparing the thermal-flow and heat transfer behavior between steady and sweeping air-only jets to ensure they are consistent with published information. Part II focuses on the mist behavior and its effect on heat transfer enhancement in the sweeping jet film cooling by adding micro-liquid droplets. An unsteady Reynolds Averaged Navier-Stokes (URANS) simulation accompanied by the k-ω SST turbulence model was used in this study. A comparison is made between steady and sweeping jets at two blowing ratios (BR = 1 and 2). The results show that the steady jet provided better film cooling performance along the centerline compared to that of the sweeping jet for both blowing ratios. However, the sweeping jet provided better and more uniform film cooling performance in the spanwise direction. Both jets experienced a significant jet-liftoff when the blowing ratio was 2. The entrainment was significant in the sweeping jet case for both blowing ratios. The sweeping jet caused an increase of 9.5% in total pressure losses compared to the steady jet. It was found that, for the sweeping jet, a pair of counter-rotating vortices is inward-rushing toward the wall in the center rather than outward-rushing as in a typical steady jet film cooling flow field. A detailed analysis is presented to understand the instantaneous vortex dynamics of the sweeping jet that leads to the inward rotating CRVP. The result shows that the pair of counter-rotating vortices is just a time-averaged image of a single vortex sweeping back and forth in the domain; it does not actually exist in real time.
Life Cycle Assessment of Innovative Concentrated Solar Power Plants Using Supercritical Carbon Dioxide Mixtures
Authors: Liao X. , Chalumeau S. , Crespi F. , Prieto C. , Lopez-Roman A. , Rodriguez de Arriba P. , Martínez N. , Sánchez D. , Paggini A. , David P. L.
Published: October 2022

The SCARABEUS project, funded by the European Commission, is currently investigating the potential gains brought about by the utilization of carbon dioxide mixtures in supercritical power cycles of Concentrated Solar Power plants, in lieu of the common Rankine cycles based on steam turbines or even pure carbon dioxide cycles. The analysis has already confirmed that it is possible to attain thermal efficiencies higher than 51% when ambient temperatures exceed 40°C, which is unheard of when conventional technology or standard CO2 technology is used. Additionally, this extraordinary performance is achieved with simpler cycle layouts, therefore with lower capital costs. The additives considered include organic and inorganic compounds which are added to the raw carbon dioxide in a variable proportion, depending on the composition of the additive and on ambient temperature. Regardless, it is important to assess whether or not there is an additional environmental advantage in terms of carbon dioxide and other potential hazards brought about by the new chemicals in the system. This is presented in this paper where the results obtained so far by the consortium for the carbon footprint from a Life Cycle perspective are discussed. Along with the assumptions and methodology, the results are compared for three reference plants: state-of-the-art CSP plant based on steam turbines, innovative CSP plant using pure supercritical CO2 technology, and the SCARABEUS concept using supercritical CO2 mixtures. The results are promising as they suggest that it is possible to reduce the carbon footprint of a 110 MWe CSP plant to be significantly less than 27kgCO2/MWh from the fifth assessment report of the Intergovernmental Panel on Climate Change (IPCC AR5).