bluuebunny/crossref_metadata_2025_split huggingface dataset:
The Effect of Surface Roughness on Efficiency of Low Pressure Turbines
Authors: Vázquez Raúl , Torre Diego 
Published: November 2013

The effect of surface roughness on efficiency of low pressure turbines (LPT) was investigated experimentally in a multistage turbine high-speed rig. The rig consisted of three stages of a state-of-the-art LPT. The stages were characterized by a very high wall-slope angle, reverse cut-off design, very high lift and very high aspect ratio airfoils.
Two sets of airfoils (both stators and rotors) were tested. The first set was made of airfoils with a roughness size of 0.7 μm Ra (25–35×10−5 ks/Cm), which was representative of LPT polished airfoils. The surface finish for the second set of airfoils was 1.8 μm Ra for blades and 2.5 μm Ra for stators (approximately 90×10−5 in terms of ks/Cm for both stators and blades). The resulting roughness of this set was representative of “as-cast” airfoils of low pressure turbines.
The airfoil geometries, velocity triangles, leading and trailing edge locations and flowpath were maintained between both sets. They were tested with the same instrumentation and at the same operating conditions with the intention of determining the isolated impact of the surface roughness on the overall efficiency. The turbine characteristics: sensitivity to speed, specific work, Reynolds number and purge flows were obtained for both sets.
The comparison of the results suggests that the efficiency and capacity of both types of airfoils exhibit the same behaviour. No significant differences in the results can be distinguished for the range of operating conditions of this study.
The results agree with previous studies of distributed roughness in turbines: the use of as-cast rough airfoils in some low pressure turbines at high altitude does not introduce additional pressure losses.
Misconceptions in Test Accuracy: Confidence Intervals vs. Statistical Power — An Engine Inlet Example
Authors: Kidman David , Stevens Craig , Remund Todd , Poulson Robert , Kitto William , Vining Charles 
Published: November 2013

The United States Air Force (USAF) is continually looking for ways to improve test and evaluation techniques to ensure systems meet military requirements. Since 1997, Air Combat Command (ACC) has been successfully using Design of Experiments (DOE) to construct and analyze operational test efforts. This paper highlights recent Air Force Materiel Command (AFMC) efforts to pursue statistically defensible test techniques to aid developmental test efforts.
Defensible testing is a statistical approach which encompasses DOE but also emphasizes the need for better test planning by insisting on understanding the system under test, by requiring clear and achievable test objectives, by ensuring system performance is measurable, and by requiring that instrumentation accuracy and uncertainty are well understood and by estimating risk in test planning and evaluations which sometimes require confidence and power calculations.
This paper highlights the Air Force Flight Test Center’s (AFFTC) first steps to improve aircraft propulsion system developmental test and evaluation (T&E) through the implementation of statistically defensible test techniques. Specifically, this paper provides insight into test accuracy. Engineers typically misuse terminology to specify test accuracy (error, precision, uncertainty, confidence intervals, etc…). This paper will present a case study in which the engineer specifies test accuracy requirements in terms of margin of error using 95-percent confidence intervals (CI’s). However, statisticians recommend using “Statistical Power” to ensure desired accuracy is achieved. Which is best to achieve test goals (CI’s or Power)? This paper will provide case study based on results from a recent aircraft engine inlet upgrade and provide a comparison between margin of error (confidence intervals) and statistical power results.
A case will be made that statistical power is more mathematically rigorous than typical margin of error approaches and should be used to ensure test accuracy requirements are met.
Wall Temperature Effects on Heat Transfer Coefficient
Authors: Maffulli Roberto , He Li 
Published: November 2013

Developments of High Pressure Turbine (HPT) blades and vanes are particularly affected by the accuracy and consistency in heat transfer prediction techniques. The conventional wisdom in heat transfer analysis considers that aerodynamics fully determines Heat Transfer Coefficient (HTC) distribution along the blade surface, so that wall temperature should have no influence on HTC. The effect of the wall temperature on the heat transfer coefficient has been rarely studied, although there have been some rather scattered correlations regarding the influence of the wall-inflow temperature ratio on HTC, largely based on simple geometry configurations. There seems to be a need to answer two related questions:
a) to what extent the conventional wisdom is valid, in particular for a transonic HPT blading?
b) what is a consistent way by which the HTC should be worked out in CFD (as well as in experiment)?
In this paper, computational analyses are carried out on the effects of wall thermal boundary condition on external HTC for a 2D Nozzle Guide Vane (NGV) profile. The study is performed by using Fluent® commercial Reynolds-Averaged Navier Stokes Equations solver. Further computations are also made using a coupled solid-fluid Conjugate Heat Transfer (CHT) method with an internally cooled blade. The present results show a strong dependence of HTC on wall temperature, far above that predicted by using the existing correlations, highlighting the importance of upstream boundary layer history on the HTC distribution. Influence of the wall temperature on the trailing edge shock position has been also observed. Based on the results found, a new method is proposed, allowing to model the HTC dependence on wall temperature with much improved accuracy as demonstrated.
Modelling of a Turbulent Jet in a Gas Crossflow
Authors: Moffat Dominic L. , Burluka Alexey A. 
Published: November 2013

The first part of this work presents a comparison of predictions obtained with several two-equation type RANS turbulence models commonly used in industry against experimental data obtained by Whitelaw et al [1]. All examined models yield a relatively poor match in the flow region very close to the wall; agreement with the measurements improves significantly when moving further away from the wall. This concerns both the internal normal stress profiles and the average velocity profiles, the latter show improved prediction of the recirculation zone area when moving further into the main stream. Downstream behaviour for both models shows an excellent match more than 6 diameters away from the jet inlet, defined as the region after which the flow essentially resumes its normal duct behaviour[1].
Expanding upon these RANS results, another series of simulations using LES modelling with the standard Smagorinsky SGS model was conducted using the same grid and compared to the RANS-based results. Although performance in the most complex flow areas was slightly improved over RANS, this was at the cost of an increase of computation time by almost a factor of 6.
The next stage involved developing a code based on the model for two-phase flow described in [2] to predict the atomisation pattern for a non-vaporising (or “cold”) flow based on the parameters of the previous simulations. This model implements transport equations for the liquid mass fraction and the average surface area per unit mass along with an equation for average density; resulting in an entirely Eulerian model which can be used to predict atomisation from first principles. Current work consists in development of additional source terms describing vaporisation in a strongly turbulent environment and further coupling with a combustion model applicable to the combustion chamber of an industrial gas turbine.
Development of an Experimental Test Facility for Investigating Mist/Air Film Cooling Application in Gas Turbine Airfoils
Authors: Wang Ting , Zhao Lei 
Published: November 2013

Film cooling is a cooling technique widely used in high-performance gas turbines to protect the turbine airfoils from being damaged by hot flue gases. Motivated by the need to further improve the turbine hot section cooling performance, a new cooling scheme, mist/air film cooling is investigated. A small amount of tiny water droplets with an average diameter about 7 μm (mist) is injected into the cooling air to enhance the cooling performance. One key feature in understanding mist cooling is the ability to capture droplet information. This paper presents the experimental facility and instrumentation of a mist/air film cooling study with both heat transfer and droplet measurements.
A wind tunnel system and test facilities are built. A Phase Doppler Particle Analyzer (PDPA) system is employed to measure the two-phase flow characteristics, including droplet size, droplet dynamics, velocity, and turbulence. Infrared camera and thermocouples are both used for temperature measurements. An extensive uncertainty analysis is performed to assist in identifying large uncertainty sources and planning for experimental procedure.
It was found during the experiment design process that resolving the mist agglomeration problem is the key in successfully generating a well-controlled mist/air mixture and reducing experimental uncertainties. The test apparatus has proven to serve the purpose well to investigate mist/air film cooling with both heat transfer and droplet measurements. Selected experimental data is presented.
Influence of Large Wake Disturbances Shed From the Combustor Wall on the Leading Edge Film Cooling
Authors: Mazzoni Cosimo Maria , Klostermeier Christian , Rosic Budimir 
Published: November 2013

The first vane leading edge film cooling is challenging because of the highest thermal load and the complex flow interaction between the hot mainstream gas and the coolant flow. This interaction varies significantly from the stagnation region to the regions of high curvature and acceleration further downstream. Additionally, in industrial gas turbines with multiple combustor chambers around the annulus the first vane leading edges may also be exposed to large wake disturbances shed from the upstream combustor walls.
The influence of these vortical structures on the first vane leading edge film cooling is numerically analysed in this paper. In order to assess the capabilities of the flow solver TBLOCK to simulate these complex interactions an experimental test case is modelled numerically. The test case is available in the open literature and consists of a cylindrical leading edge and two rows of film cooling holes representative of industrial practice. A LES turbulence modelling strategy with WALE sub-grid scale (SGS) model is applied and compared against experimental results. Based on this validation it is decided to analyse also the wake–leading edge interaction, dominated by large scale unsteady vortical structures, using the same WALE sub-grid scale LES model. The initial flow domain with the cylindrical leading edge and cooling holes is extended to incorporate the effect of the combustor wall, which is modelled as a flat plate with a square trailing edge. The location and the size of the plate are scaled to be representative of industrial practice: the plate is located upstream from the leading edge at a distance twice the leading edge diameter, and the thickness of the plate is one half of the leading edge diameter. Two different clockwise positions of the vertical combustor wall model were investigated and compared with the datum configuration: the former where the axis of the plate and the leading edge are aligned (central wake location), the latter with the combustor wall circumferentially shifted up by a quarter of the leading edge diameter (circumferentially shifted wake location).
Numerical predictions show that the shed vortices from the combustor wall trailing edge have a highly detrimental effect on the leading edge film cooling by periodically removing the coolant flow from the leading edge surface. This results in an increased unsteady thermal load. These negative effects are less significant in the case of circumferentially shifted wake, due to the combined action of both shed vortices.
Conjugate Heat Transfer Modeling of a Film-Cooled, Flat-Plate Test Specimen in a Gas Turbine Aerothermal Test Facility
Authors: Sidwell T. G. , Lawson S. A. , Straub D. L. , Casleton K. H. , Beer S. 
Published: November 2013

The aerothermal test facility at the National Energy Technology Laboratory (NETL) provides experimental data at realistic gas turbine conditions to enable the development of advanced film cooling strategies for future gas turbine components. To complement ongoing experimental studies, Fluent computational fluid dynamics (CFD) models have been developed to provide a framework for comparison of cooling strategies and to provide fundamental understanding of the fluid dynamic and conjugate heat transfer (CHT) processes occurring in the experiments.
The results of a parametric study of the effects of mesh density, near-wall refinement, wall treatment, turbulence model and gradient discretization order on the CHT predictions are presented, and the simulation results are compared to experimental data. A flat plate test specimen with a single row of laidback fan-shaped film cooling holes was modeled at a process pressure of 3 bar, a process gas flow rate (m) of 0.325 kg/s (Re ≈ 100,000) and a blowing ratio (M) of 2.75. Three polyhedral mesh cases and three turbulence models (Realizable k-ε, SST k-ω and RSM Stress-ω) were implemented with enhanced wall treatment (EWT) and 1st-order and 2nd-order gradient discretization.
The results show that the choice of turbulence model will have little effect on the results when utilizing the finest mesh case and 2nd-order discretization. It was also shown that the SST k-ω turbulence model cases showed minimal mesh sensitivity with 2nd-order discretization, while the Re k-ε turbulence model cases were more sensitive to mesh density and near-wall refinement. The results thus indicate that the SST k-ω turbulence model can predict the convective heat transfer adequately with a relatively coarse mesh, which will save computational resources for later inclusion of radiative heat transfer effects to provide comprehensive CHT predictions.
Experimental and Numerical Investigations of Flowpath Profiling on Secondary Flow Losses in a Turbine Control Stage
Authors: Moser Nils , Steinhoff Peter , Joos Franz 
Published: November 2013

The numerically and experimentally investigated industrial steam turbine control stage is derived from a real design. Due to the production process and costs of the guide vanes for control stages of steam turbines the flowpath profiling is rotationally symmetric. However the combination of the two-dimensional shroud contour and the flow deflection in the guide vane results in a fully three-dimensional end wall contour having a strong influence on the secondary flow features in the turbine control stage.
To obtain an improved profile for the nozzle shroud the reduction of the total pressure loss over the guide vanes is taken as an optimization criterion. The three-dimensional contour generates a diffuser flowpath between the suction and the pressure side of two guide vanes perpendicular to the main flow direction. This diffuser geometry affects the pressure distribution over the guide vane and therefore the formation mechanisms of secondary flows. For the experimental and numerical investigations a baseline shroud design and two additional profiled contours are analyzed in detail.
The control stage test rig is operated with air and is capable to represent a wide range of operating conditions. The measurements show a considerable increase of the stage efficiency and power output. The effect of the flowpath profiling on the pressure distribution over the guide vane is clearly proved.
A Particle Image Velocimetry-Based Investigation of the Flow Field in an Oblique Jet Impingement Configuration
Authors: Schulz Sebastian , Schueren Simon , von Wolfersdorf Jens 
Published: November 2013

Impinging jets have become an indispensable measure for cooling applications in gas turbine technology. The present study seeks to unfold the flow field dynamics inside an enigine-relevant cooling passage of trapezoidal cross-section. The investigated geometry produces a highly complex flow field which was investigated employing Particle Image Velocimetry (PIV). The experiments were accompanied by numerical simulations solving the RANS equations with FLUENT using the low-Re k-ω-SST turbulence model. Additionally, time-resolved pressure measurements were performed utilizing Kulite pressure transducers. The spectral analysis of the transient pressure signal in conjunction with a Proper Orthogonal Decomposition (POD) analysis of the PIV data allows a detailed insight into the effects of geometric constraints on the fluid dynamic processes inside the geometry.
The results are presented for a jet Reynolds number of 45,000 and display a qualitatively fair agreement between the experiments and numerical simulations. Nevertheless, the simulations predict flow features in particular regions of the geometry that are absent in the experiments. Despite the lack of conspicuous high energy modes, the flow was well suited for a POD analysis. Depending on the considered PIV plane, it could be shown that up to 25% of the flow field’s total turbulent energy are contained in the first ten POD modes. Additionally, using the first 20 to 60 POD modes sufficed to reconstruct the flow fields with its governing features.
Influence of Pressure and Steam Dilution on NOx and CO Emissions in a Premixed Natural Gas Flame
Authors: Göke Sebastian , Schimek Sebastian , Terhaar Steffen , Reichel Thoralf , Göckeler Katharina , Krüger Oliver , Fleck Julia , Griebel Peter , Paschereit Christian Oliver 
Published: November 2013

In the current study, the influence of pressure and steam on the emission formation in a premixed natural gas flame is investigated at pressures between 1.5 bar and 9 bar.
A premixed, swirl-stabilized combustor is developed that provides a stable flame up to very high steam contents. Combustion tests are conducted at different pressure levels for equivalence ratios from lean blowout to near-stoichiometric conditions and steam-to-air mass ratios from 0% to 25%.
A reactor network is developed to model the combustion process. The simulation results match the measured NOx and CO concentrations very well for all operating conditions. The reactor network is used for a detailed investigation of the influence of steam and pressure on the NOx formation pathways.
In the experiments, adding 20% steam reduces NOx and CO emissions to below 10 ppm at all tested pressures up to near-stoichiometric conditions. Pressure scaling laws are derived: CO changes with a pressure exponent of approximately −0.5 that is not noticeably affected by the steam. For the NOx emissions, the exponent increases with equivalence ratio from 0.1 to 0.65 at dry conditions. At a steam-to-air mass ratio of 20%, the NOx pressure exponent is reduced to −0.1 to +0.25.
The numerical analysis reveals that steam has a strong effect on the combustion chemistry. The reduction in NOx emissions is mainly caused by lower concentrations of atomic oxygen at steam-diluted conditions, constraining the thermal pathway.
Effect of Variable Properties Within a Boundary Layer With Large Freestream to Wall Temperature Differences
Authors: Greiner Nathan J. , Polanka Marc D. , Robertson Jacob R. , Rutledge James L. 
Published: November 2013

Modern aviation combustors run at high fuel-air ratios to achieve high turbine inlet temperatures and higher turbine efficiencies. To maximize turbine durability in such extreme temperatures, the blades are fitted with film cooling schemes to form a layer of cool air between the blade and the hot core flow. Two terms that are utilized to evaluate a cooling scheme are the heat transfer coefficient (h) and the local driving temperature, namely, the adiabatic wall temperature (Taw). The literature presents a method for calculating these two parameters by assuming the heat flux (q) is proportional to the difference in freestream and wall temperatures (T∞ − Tw). Several researchers have shown the viability of this approach by altering the wall temperature over a finite range in low temperature environment. A linear trend ensues where the slope is h and the q = 0 intercept is adiabatic wall temperature. This technique has proven valuable since constant h is known to be a valid assumption for constant property flow.
The current study explores the validity of this assumption by analytically predicting and experimentally measuring the h and q at high T∞ and low Tw characteristic of a modern combustor. Both a reference temperature method and temperature ratio method were applied to model the effects of variable properties within the boundary layer. To explore the linearity of the heat transfer with driving temperature, the analysis determined the apparent h and Taw which would be measured over small ranges of Tw by the linear method discussed in the literature. This study shows that, over large Tw ranges, property variations play a significant role. It is also shown that the linear trend technique is valid even at high temperature conditions but only when used in small temperature ranges. Finally, this investigation shows that the apparent Taw used in the linear convective heat transfer assumption is a valid driving temperature over small ranges of Tw but cannot always be interpreted literally as the temperature where q(Taw) = 0.
Characterisation of Lean Burn Module Air Blast Pilot Injector With Laser Techniques
Authors: Meier U. , Freitag S. , Heinze J. , Lange L. , Magens E. , Schroll M. , Willert C. , Hassa C. , Bagchi I. K. , Lazik W. , Whiteman M. 
Published: November 2013

For lean burn combustor development in low emission aero-engines, the pilot stage of the fuel injector plays a key role with respect to stability, operability, NOx emissions, and smoke production. Therefore it is of considerable interest to characterize the pilot module in terms of pilot zone mixing, fuel placement, flow field and interaction with the main stage.
This contribution focusses on the investigation of soot formation during pilot-only operation. Optical test methods were applied in an optically accessible single sector rig at engine idle conditions. Using planar laser-induced incandescence (LII), the distribution of soot and its dependence on air/fuel ratio, as well as geometric injector parameters, was studied. The data shows that below a certain air/fuel ratio, an increase of soot production occurs. This is in agreement with smoke number measurements in a standard single sector flame tube rig without optical access. Reaction zones were identified using chemiluminescence of OH radicals. In addition, the injector flow field was investigated with PIV. A hypothesis regarding the mechanism of pilot smoke formation was made based on these findings. This along with further investigations will form the basis for developing strategies for smoke improvement at elevated pilot only conditions.