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Process capability indices (PCIs) are major tools in geometric dimensioning and tolerancing (GD&T) for quantifying the production quality, monitoring production, or prioritizing projects. Initially, PCIs were constructed for studying each characteristic of the process independently. Then, they have been extended to analyze several dependent characteristics simultaneously. Nowadays, with the increasing complexity of the production parts, for example, in aircraft engines, the conformity of one part may rely on the conformity of hundreds of characteristics. Also, because those characteristics are dependent, it may be misleading to make decisions based only on univariate PCIs. However, classical multivariate PCIs in the literature do not allow treating such amount of data efficiently, unless assuming Gaussian distribution, which may be false. Regarding those issues, we advocate for PCI based on some transformation of the conformity rates. This presents the advantage of being free from distributional assumptions, such as the Gaussian distribution. In addition, it has direct interpretation, allowing it to compare different processes. To estimate the PCIs of parts with hundreds of characteristics, we propose to use Vine Copulas. This is a very flexible class of models, which gives precise estimation even in high dimension. From an industrial perspective, the computation of the estimator can be costly. To answer this point, one explains how to compute a lower bound of the proposed PCI, which is faster to calculate. We illustrate our method adaptability with simulations under Gaussian and non-Gaussian distributions. We apply it to compare the production of fan blades of two different factories. [ABSTRACT FROM AUTHOR]

Complicated systems made of multiple components are known to be difficult to model, considering their solutions can change dramatically even with the slightest variations in conditions. Aircraft engines contain such complicated systems, and some components in aircraft engines' turbines can cause significant changes in the system's overall response. Hence, this study is focused on investigating the behavior of a turbine blade of an aircraft engine and the effects of the contact between the blade and the seal wire on the dynamics of the blade-disk system. The investigation is performed via various numerical simulations in time and frequency domains. One sector of the bladed disk is modeled using the finite element method with the lock plate and the seal wire imposing cyclic symmetry boundary conditions in the static, modal, and frequency domain forced response analyses. In time domain analyses, the cyclic symmetry is replaced with simplified displacement restricting boundary conditions. The time domain analysis contains steady-state forced responses of the system. The results show that contact with the seal wire is not a major source of nonlinearity and damping. The contacts with the lock plate contribute more to the vibration damping than the seal wire. However, compared to the contacts at the root of the blade, both components remain less significant with regard to frictional damping and nonlinearity. [ABSTRACT FROM AUTHOR]

Usually, the process capability indices (PCIs) Cp and Cpk are used to quantify the production quality of Geometric Dimensioning & Tolerancing (GD&T). For the most popular PCIs, a strong assumption about normality is required. Many previous works provide information about misleading conclusions when we use usual PCIs without any verification of the dataset normality. In recent industrial developments, many processes increasingly complicate, and GD&T are mostly geometrical specifications. For that reason, most characteristics are not normally distributed. There are two issues when calculating PCIs with nonnormal distributions. First, we have to define PCIs that can deal with all distributions. Second, we need to determine the distribution of the dataset. To do so, we give a short review of the previous work about calculating PCIs with nonnormal distributions. Following that first part, we propose our method for nonnormal distributions. This method uses new PCIs based on the quantile of a standard normal distribution. Then, we consider an estimation method based on a Student mixture model. This allows handling nonnormal distributions, potentially with outliers, and mixture distributions, which are often treated as two different fields and are common in industrial datasets. Finally, we propose some use cases to benchmark our new PCIs and Student mixture model for aircraft engines GD&T applications. [ABSTRACT FROM AUTHOR]

Due to its ability to maximize thrust vectoring performance, the bypass dual throat nozzle (BDTN) has an advantage over other fluidic vectoring controls. In this study, numerical simulation is performed to analyze the flow characteristics and performance parameters of an aircraft engine with three different nozzle configurations. The nozzle of a representative engine, i.e., an F100 engine was selected as a model geometry to test the efficiency obtained by BDTN. The present investigation has shown that implementing a bypass channel on a real geometry nozzle has no significant effects on thrust vectoring performance in vectored mode. Although the real geometry scheme has a higher thrust and a discharge coefficient, the smaller cavity length resulted in lower vectoring angles. Modifying the real geometry nozzle according to the BDTN configuration significantly improved the thrust vectoring performance. However, the V-shaped bypass passage flow in the modified geometry scheme imposed un-necessary total pressure losses in the nozzle. A geometry scheme that utilized an arc-shaped rather than a V-shaped bypass passage is considered in this research and found to present minimize pressure losses. A total increase of 2% and 3.5% is hereby reported, for thrust and discharge coefficients, respectively. A decrease of 6% is reported in the thrust vectoring angle under an improved geometrical scheme. Out of three geometries, the real geometry scheme reported negligible thrust vectoring performance, while modified and improved geometry schemes indicated improved thrust vectoring performance without substantially changing the engine states. [ABSTRACT FROM AUTHOR]

Contemporary squeeze film dampers (SFDs) in air-breathing engines are short in length to limit weight and part count and lubricated with a low feed pressure to reduce oil storage and pumping power. In SFDs, O-rings (ORs) restrict side leakage and increase the viscous damping while a adding a modest centering stiffness. Continuing a long-term project characterizing SFDs for aircraft engines and extending the original work (San Andrés and Rodríguez, 2021, "On the Experimental Dynamic Force Performance of a Squeeze Film Damper Supplied Through a Check Valve and Sealed With O-Rings," ASME J. Eng. Gas Turb. Power, 143(11), p. 111011) the paper details measurements of the forced performance of an OR sealed damper (OR-SFD) with diameter D = 127 mm, land length L = 0.2 D, and radial clearance c = 0.0022 D. Lubricant ISO VG 2 supplied at 0.69 bar(g) fills an upstream oil plenum and flows into the middle of the land through a single orifice configured with a check valve. Measurements of applied single-frequency dynamic loads, along with the ensuing damper displacements and accelerations serve to identify the parameters of the test structure, ORs, and SFD. The tests encompass centered whirl motions with amplitude r = 0.05--0.45c, and a range of whirl frequencies, ω = 10--130 Hz. Note the squeeze velocity vs = rω reaches 102 mm/s. The ORs force coefficients are nearly invariant with frequency but do depend on the orbit amplitude. The ORs' stiffness (KOR) decreases by 75% as the motion amplitude increases, r → 0.45c, likely due to the large elastic deformations and slow recovery of the ORs material. For small amplitude motions (r/c = 0.05 and 0.10), the ORs damping coefficient (COR) is ~10% of the overall coefficient for the lubricated system (CL), while for r/c > 0.25, COR ~ 0.03CL. For small amplitudes of whirl (r ≤ 0.25c), the SFD experimental viscous damping (CSFD) and added mass (MSFD) coefficients, identified over a shorter frequency range (vs<30 mm/s) equal theoretical magnitudes for a fully sealed damper. As the orbit size grows to r = 0.45c MSFD drops nearly by 75% and CSFD decreases by ~40% due to the onset of both lubricant cavitation and air ingestion occurring for squeeze velocities vs ≥ 24.5 mm/s, as also seen in the recorded dynamic pressures and video recordings of a bubbly mixture leaving the damper. A comprehensive flow model predicts CSFD and MSFD about 8% and 12% larger than the experimentally identified coefficients. A novel approach enables the estimation of the gas volume fraction (GVF) generated in the damper which rapidly increases as vs grows. The simple procedure draws into a deflated balloon the material contents in the squeeze film, weighs the sample, and identifies its volume to produce an estimate of the GVF. The procedure to quantify the GFV will assist in the validation of predictive tools. [ABSTRACT FROM AUTHOR]

To deal with class imbalance learning (CIL) problems, a novel algorithm is proposed based on kernel extreme learning machine (KELM), named KELM-CIL. To solve it, two algorithms are developed from the dual and primal spaces, respectively, thus yielding D-KELM-CIL and P-KELM-CIL. However, both D-KELM-CIL and P-KELM-CIL are not sparse algorithms. Hence, a sparse strategy based on Cholesky factorization is utilized to realize their sparseness, producing CD-KELM-CIL and CP-KELM-CIL. For large-size problems, a probabilistic trick is applied to accelerate them further, hence obtaining PCD-KELM-CIL and PCP-KELM-CIL. To test the effectiveness and efficacy of the proposed algorithms, experiments on benchmark datasets are carried out. When the proposed algorithms are applied to fault detection of aircraft engine, they show good generalization performance and real-time performance, especially for CP-KELM-CIL and PCP-KELM-CIL, which indicates that they can be developed as candidate techniques for fault detection of aircraft engine. [ABSTRACT FROM AUTHOR]

In aircraft engine control, replacing linear regulators by sliding mode control (SMC) regulators is considered as an effective approach to reducing the conservatism in the traditional treatment for limit protection. However, most of the relevant studies are based on linear descriptions, which cannot represent the nonlinear systems directly due to their limited valid range. Even if gain scheduling techniques are employed, the stability of the nonlinear systems cannot be theoretically guaranteed. In this paper, a sliding mode strategy for a class of uncertain linear parameter varying (LPV) systems is studied. LPV descriptions are applied to extend the valid range of the linear models covering the entire operation envelope with guaranteed performance and stability. The mismatch between LPV and the real systems is considered as uncertainties. With a sliding surface defined by the tracking errors, system properties on the surface are proved to be satisfactory. After that, a reaching law is designed to ensure global invariance of SMC. Based on a reliable model turbofan, simulation results show that the SMC method can fully exploit the limit margin and, compared to the traditional proportional-integral-derivative (PID) control, has a faster response. In addition, stability and effectiveness of the proposed method are verified in a temperature protection case. [ABSTRACT FROM AUTHOR]

Removing noise from health signals is critical in gas path diagnostics of aircraft engines. An efficient noise filtering/denoising method should remove noise without using future data points, preserve important changes, and promote accurate diagnostics without time delay. Machine learning (ML)-based methods are promising for high fidelity, accuracy, and computational efficiency under the motivation of Intelligent Engines. However, previous ML-based denoising methods are rarely applied in actual engineering practice because they cannot accommodate time series and cannot effectively capture important changes or are limited by the time delay problem. This paper proposes a convolutional neural network denoising auto-encoder (CNN-DAE) method to build a denoising auto-encoder structure. In this structure, a convolutional operation is used to accommodate time series, and causal convolution is introduced to solve the problem of using future data points. The proposed denoising method is evaluated against NASA's propulsion diagnostic method evaluation strategy (prodimes) software. It has been proved that the proposed method can accommodate time series, remove noise for improved denoising accuracy, and preserve the important changes for enhanced diagnostic information. NASA's blind test case results show that kappa coefficient of a common diagnostic method using the processed data is 0.731 and is at least 0.046 higher than the other diagnostic methods in the open literature. Processing health signals using the proposed method would significantly promote accurate diagnostics without time delay. The proposed method could support intelligent condition monitoring systems by exploiting historical information for improved denoising and diagnostic performance. [ABSTRACT FROM AUTHOR]

A novel theoretical model of the internal flow field in multistage axial compressors based on an eigenvalue approach is developed, in order to predict the onset of acoustic resonance in aircraft engines. Using an example high-speed four-stage compressor, it is shown that one of the resultant frequencies is in excellent agreement with the experimental data in terms of acoustic resonance. On the basis of the computed natural frequency of the whole compression system and the measured spanwise distribution of static pressure, the location of the acoustic excitation source can be found in the third stage. Unsteady flow simulations of the full annulus of this stage reveal two criteria for acoustic excitation at the rotor-blade tip, reversed flow near the suction surface and flow impingement on the pressure surface. Additionally, a fast Fourier transform of the unsteady pressure field at the upper rotor-blade span verifies the existence of the computed unstable frequency of the oscillating tip leakage flow. Using this novel theory, which combines a theoretical calculation of flow-instability frequency of the global system with the computational simulation of a single stage, the onset mechanism and location of the excitation source of acoustic resonance in multistage turbomachinery can be explained at acceptable computational cost. [ABSTRACT FROM AUTHOR]

Numerical optimizations are very useful in liner designs for low-noise aeroengines. Although modern computational tools are already very efficient for a single aeroengine noise propagation simulation run, the prohibitively high computational cost of a broadband liner optimization process which requires hundreds of thousands of runs renders these tools unsuitable for such task. To enable rapid optimization using a desktop computer, an efficient analytical solver based on the Wiener-Hopf method is proposed in the current study. Although a Wiener-Hopf-based solver can produce predictions very quickly (order of a second), it usually assumes an idealized straight duct configuration with a uniform background flow that makes it arguable for practical applications. In the current study, we employ the Wiener-Hopf method in our solver to produce an optimized liner design for a semi-infinite annular duct setup and compare its noise-reduction effect with an optimized liner designed by the direct application of a numerical finite element solver for a practical aeroengine intake configuration with an inhomogeneous background flow. The near-identical near- and far-field solutions by the Wiener-Hopf-based method and the finite element solvers clearly demonstrate the accuracy and high efficiency of the proposed optimization strategy. Therefore, the current Wiener-Hopf solver is highly effective for liner optimizations with practical setups and is very useful to the preliminary design process of low-noise aeroengines. [ABSTRACT FROM AUTHOR]

The steady improvement of aircraft engine performance has led toward more compact engine cores with increased structural loads. Compact single-stage high-pressure turbines allow high power extraction, operating in the low supersonic range. The shock waves formed at the airfoil trailing edge contribute substantially to turbine losses, mainly due to the shock-boundary layer interactions as well as high-frequency forces on the rotor. We propose to control the vane trailing edge shock interaction with the downstream rotor, using a pulsating vane-trailing-edge-coolant at the rotor passing frequency. A linear cascade of transonic vanes was investigated at different Mach numbers, ranging from subsonic to supersonic regimes (0.8, 1.1) at two engine representative Reynolds numbers (4 × 106 and 6 × 106). The steady and unsteady heat flux was retrieved using thin-film two-layered gauges. The complexity of the tests required the development of an original heat transfer postprocessing approach. In a single test, monitoring the heat flux data and the wall temperature we obtained the adiabatic wall temperature and the convective heat transfer coefficient. The right-running trailing edge shock wave impacts on the neighboring vane suction side. The impact of the shock wave on the boundary layer creates a separation bubble, which is very sensitive to the intensity and angle of the shock wave. Increasing the coolant blowing rate induces the shock to be less oblique, moving the separation bubble upstream. A similar effect is caused by the pulsations of the coolant. [ABSTRACT FROM AUTHOR]

Airborne particles ingested in aircraft engines deposit on compressor blading and end walls. Aerodynamic surfaces degrade on a microscopic and macroscopic scale. Blade row, compressor, and engine performance deteriorate. Optimization of maintenance scheduling to mitigate these effects requires modeling of the deterioration process. This work provides a deterioration model on blade row level and the experimental validation of this model in a newly designed deposition test rig. When reviewing previously published work, a clear focus on deposition effects in industrial gas turbines becomes evident. The present work focuses on quantifying magnitudes and timescales of deposition effects in aircraft engines and the adaptation of the generalized Kern and Seaton deposition model for application in axial compressor blade rows. The test rig's cascade was designed to be representative of aircraft engine compressor blading. The cascade was exposed to an accelerated deposition process. Reproducible deposition patterns were identified. Results showed an asymptotic progression of blade row performance deterioration. A significant increase in total pressure loss and decrease in static pressure rise were measured. Application of the validated model using existing particle concentration and flight cycle data showed that more than 95% of the performance deterioration due to deposition occurs within the first 1000 flight cycles. [ABSTRACT FROM AUTHOR]

Aircraft engines ingest airborne particulate matter, such as sand, dirt, and volcanic ash, into their core. The ingested particulate is transported by the secondary flow circuits via compressor bleeds to the high pressure turbine and may deposit resulting in turbine fouling and loss of cooling effectiveness. Prior publications focused on particulate deposition and sand erosion patterns in a single stage of a compressor or turbine. This work addresses the migration of ingested particulate through the high pressure compressor (HPC) and bleed systems. This paper describes a 3D CFD methodology for tracking particles along a multistage axial compressor and presents particulate ingestion analysis for a high pressure compressor section. The commercial CFD multiphase solver ANSYS CFX® has been used for flow and particulate simulations. Particle diameters of 20, 40, and 60µm are analyzed. Particle trajectories and radial particulate profiles are compared for these particle diameters. The analysis demonstrates how the compressor centrifuges the particles radially toward the compressor case as they travel through the compressor; the larger diameter particles being more significantly affected. Nonspherical particles experience more drag as compared to spherical particles, and hence a qualitative comparison between spherical and nonspherical particles is shown. [ABSTRACT FROM AUTHOR]

Nonlinear control of aircraft engines has attracted much attention in consideration of the inherent nonlinearity of the engine dynamics. Most of the nonlinear design techniques, however, require the information from the rotational speeds of both high-pressure compressor and fan. This is not desirable from engine health management perspective, and this paper proposes a single sensor measurement and single actuator control approach. The proposed method can provide fast regulation of engine speed in a finite-time in comparison with conventional infinite time stability. Important results are obtained on both controller design and disturbance tolerance. Numerical examples are provided for validation of the proposed finite-time controller, demonstrating fast regulation property and remarkable disturbance tolerance capability. [ABSTRACT FROM AUTHOR]

This paper presents analytical techniques for aiding system designers in making aircraft engine health management sensor selection decisions. The presented techniques, which are based on linear estimation and probability theory, are tailored for gas turbine engine peiformance estimation and gas path fault diagnostics applications. They enable quantification of the peiformance estimation and diagnostic accuracy offered by different candidate sensor suites. For peiformance estimation, sensor selection metrics are presented for two types of estimators including a Kalman filter and a maximum a posteriori (MAP) estimator. For each type of peiformance estimator, sensor selection is based on minimizing the theoretical sum of squared estimation errors (SSEE) in health parameters representing peiformance deterioration in the major rotating modules of the engine. For gas path fault diagnostics, the sensor selection metric is set up to maximize correct classification rate (CCR) for a diagnostic strategy that performs fault classification by identifying the fault type that most closely matches the observed measurement signature in a weighted least squares sense. Results from the application of the sensor selection metrics to a linear engine model are presented and discussed. Given a baseline sensor suite and a candidate list of optional sensors, an exhaustive search is performed to determine the optimal sensor suites for performance estimation and fault diagnostics. For any given sensor suite, Monte Carlo simulation results are found to exhibit good agreement with theoretical predictions of estimation and diagnostic accuracies. [ABSTRACT FROM AUTHOR]

In modern aero-engines, the lubrication system holds a key role due to the demand for high reliability standards. An aero-engine bearing chamber contains components like bearings and gears. Oil is used for lubrication and for heat removal. In order to retain the oil in a bearing chamber, pressurized seals are used. These are pressurized using air from the compressor. In order to avoid overpressurization of the bearing chamber, air/oil passages are provided in the bearing chamber. At the top, a vent pipe discharges most of the sealing air and at the bottom, a scavenge pipe is used for discharging the oil by means of a pump (scavenge pump). The scavenge pipe is setup in most cases by tubes of circular or noncircular cross sections. When the scavenge pipe has to be routed in a way that sharp bends or elbows are unavoidable, flexible (corrugated) pipes can be used. Because of the corrugation, considerable flow resistance with high-pressure drop can result. This may cause overpressurization of the bearing compartment with oil loss into the turbomachinery with possibility of ignition, coking (carbon formation), or contamination of the aircraft's air conditioning system. It is therefore important for the designer to be capable to predict the system's pressure balance behavior. A real engine bearing chamber sealed by brush seals was used for generating different air/oil mixtures thus corresponding to different engine operating conditions. The mixtures were discharged through a scavenge pipe which was partly setup by corrugated tubes. Instead of a mechanical pump, an ejector was used for evacuating the bearing chamber. An extensive survey covering the existing technical literature on corrugated tube pressure drop was performed and is presented in this paper. The survey has covered both single-phase and multiphase flows. Existing methods were checked against the test results. The method which was most accurately predicting lean air test results from the rig was benchmarked and was used as the basis for extending into a two-phase flow pressure drop correlation by applying two-phase flow multiplier techniques similar to Lockhart and Martinelli. Comparisons of the new two-phase flow pressure drop correlation with an existing correlation by Shannak are presented for mixtures like air/oil, air/water, air/diesel, and air/kerosene. Finally, numerical analysis results using ANSYS CFX VERSION 15 are presented. [ABSTRACT FROM AUTHOR]

During the last decades, turbo machine efficiency was considerably increased by using more efficient seals. Brush seals, as a compliant contacting filament seal, have become an attractive alternative to conventional labyrinth seals in the field of aircraft engines as well as in stationary gas and steam turbines. The aim of today's research related to brush seals is to understand the characteristics and their connections, in order to be able to make performance predictions, and to ensure the reliability over a defined operating period. The wear behavior is essentially influenced by frictional contacts at the seal-to-rotor interface during operation. For realistic investigations with representative circumferential velocities, the TU Braunschweig, Germany, operates a specially developed steam test rig which enables endurance investigations under varying operating steam conditions up to 50 bar and 450°C. Wear measurements and the determination of seal performance characteristics, such as blow down and bristle stiffness, are enabled by an additional test facility, using pressurized cold air up to 8 bar as a working fluid. This work presents the chronological wear development on both rotor and seal sides in a steam test lasting 25 days or 11 days, respectively. Interruptions after stationary and transient intervals were made in order to investigate the degree of wear. Two different seal arrangements, a single tandem seal, and a two-stage single seal arrangement, using different seal elements were considered. Besides a continuous wear development, the results clearly show that the abrasive wear of the brush seal and rotor is mainly caused by transient test operations, particularly by enforced contacts during shaft excursions. Despite the increasing wear to the brushes, all seals have shown a functioning radial-adaptive behavior over the whole test duration with a sustained seal performance. Thereby, it could be shown that the two-stage arrangement displays a load shift during transients, leading to a balanced loading and unloading status for the two single brush seals. From load sharing, and in comparison with the wear data of the tandem seal arrangement, it can be derived that the two-stage seal is less prone to wear. However, the tandem seal arrangement, bearing the higher pressure difference within one configuration, shows a superior sealing performance under constant load, i.e., under stationary conditions. [ABSTRACT FROM AUTHOR]

International standards regarding polluting emissions from civil aircraft engines are becoming gradually even more stringent. Nowadays, the most prominent way to meet the target of reducing NOx emissions in modern aero-engine combustors is represented by lean-burn technology. Swirl injectors are usually employed to provide the dominant flame stabilization mechanism coupled to high-efficiency fuel atomization solutions. These systems generate very complex flow structures, such as recirculations, vortex breakdown, and processing vortex core, which affect the distribution and therefore the estimation of heat loads on the gas side of the liner as well as the interaction with the cooling system flows. The main purpose of the present work is to provide detailed measurements of heat transfer coefficient (HTC) on the gas side of a scaled combustor liner highlighting the impact of the cooling flows injected through a slot system and an effusion array. Furthermore, for a deeper understanding of the interaction phenomena between gas and cooling flows, a standard two-dimensional (2D) particle image velocimetry (PIV) technique has been employed to characterize the combustor flow field. The experimental arrangement has been developed within EU project LEMCOTEC and consists of a nonreactive three sectors planar rig installed in an open-loop wind tunnel. Three swirlers, replicating the real geometry of a GE Avio partially evaporated and rapid mixing (PERM) injector technology, are used to achieve representative swirled flow conditions in the test section. The effusion geometry is composed by a staggered array of 1236 circular holes with an inclination of 30 deg, while the slot exit has a constant height of 5 mm. The experimental campaign has been carried out using a thermochromic liquid crystals (TLCs) steady-state technique with a thin Inconel heating foil and imposing several cooling flow conditions in terms of slot coolant consumption and effusion pressure drop. A data reduction procedure has been developed to take into account the nonuniform heat generation and the heat loss across the liner plate. Results in terms of 2D maps and averaged distributions of HTC have been supported by flow field measurements with 2D PIV technique focussed on the corner recirculation region. [ABSTRACT FROM AUTHOR]

Flow and heat transfer in an aero-engine compressor disk cavity with radial inflow has been studied using computational fluid dynamics (CFD), large eddy simulation (LES), and coupled fluid! solid modeling. Standalone CFD investigations were conducted using a set of popular turbulence models along with 0.2 deg axisymmetric and a 22.5 deg discrete sector CFD models. The overall agreement between the CFD predictions is good, and solutions are comparable to an established integral method solution in the major part of the cavity. The LES simulation demonstrates that flow unsteadiness in the cavity due to the unstable thermal stratification is largely suppressed by the radial inflow. Steady flow CFD modeling using the axisymmetric sector model and the Spalart-Allmaras turbulence model was coupled with a finite element (FE) thermal model of the rotating cavity. Good agreement was obtained between the coupled solution and rig test data in terms of metal temperature. Analysis confirms that using a small radial bleed flow in compressor cavities is effective in reducing thermal response times for the compressor disks and that this could be applied in management of compressor blade clearance. [ABSTRACT FROM AUTHOR]

Squeeze film dampers (SFDs) are effective to ameliorate shaft vibration amplitudes and to suppress instabilities in rotor-bearing systems. Compact aero jet engines implement ultra-short length SFDs (LID ⩽0.2) to satisfy stringent weight and space demands with low parts count. This paper describes a test campaign to identify the dynamic forced response of an open ends SFD (L = 25.4 mm and D = 125.7 mm), single film land, and oil fed through three holes (120deg apart), operating with similar conditions as in an aircraft engine. Two journals make for two SFD films with clearances cA = 0.129 mm and cb = 0 .2 5 4 mm (small and large). The total oil-wetted length equals Lto, = 36.8 mm that includes deep end grooves, width and depth = 2.5 × 3.8 mm, for installation of end seals. In the current experiments, the end seals are not in place. A hydraulic static loader pulls the bearing cartridge (BC) to a preset static eccentricity (es), and two electromagnetic shakers excite the BC with single frequency loads to create circular orbits, centered and off-centered, over a prescribed frequency range ω = 10-100 Hz. The whirl amplitudes range from r = 0.05cA-0.6cA and r = 0.15cB-0.75cB while the static eccentricity increases to es = 0.5cA and es = 0.75cB, respectively. Comparisons of force coefficients between the two identical dampers with differing clearances show that the small clearance damper (cA) provides ~4 times more damping and .8 times the inertia coefficients than the damper with large clearance (cB). The test results demonstrate damping scales with ~1/c3 and inertia with ~1lc, as theory also showed. Analysis of the measured film land pressures evidence that the deep end grooves contribute to the generation of dynamic pressures enhancing the dynamic forced response of the test SFDs. A thin film flow model with an effective groove depth delivers predictions that closely match the test damping and inertia coefficients. Other predictions, based on the short length bearing model, use an effective length Leff~1.17L to deliver damping coefficients 15% larger than the experimental results; however, inertia coefficients are 1/2 of the identified magnitudes. The experiments and analysis complement earlier experimental work conducted with centrally grooved SFDs. [ABSTRACT FROM AUTHOR]

Testing and simulation of aero-engine spectra with dwell times are reported in this paper. The modeling concept used is built on linear elastic fracture mechanics (LEFM) and provides a history-dependent evolution description of dwell damage and its interaction with cyclic load. The simulations have been carried out for three spectra: (1) cyclic loads, (2) combined sustained load and cyclic loads, and (3) slow load ramps and cyclic loads, all for surface cracks at 550 °C for Inconel 718. All simulations show reasonable good agreement with experimental results. Prediction of multiple tests of several batches is also provided to show statistical scatter. [ABSTRACT FROM AUTHOR]

Online line-of-sight (LOS) pyrometer is used on certain jet engines for diagnosis and control functions such as hot-blade detection, high-temperature limiting, and conditionbased monitoring. Hot particulate bursts generated from jet engine combustor at certain running conditions lead to intermittent high-voltage signal outputs from the LOS pyrometer which is ultimately used by the onboard digital engine controller (DEC). To study the nature of hot particulates and enable LOS pyrometer functioning under burst conditions, a multicolor pyrometry (MCP) system was developed under DARPA funded program and tested on an aircraft jet engine. Soot particles generated as byproduct of combustion under certain conditions was identified as the root cause for the signal burst in a previous study. The apparent emissivity was then used to remove burst signals. In current study, the physics based filter with MCP algorithm using apparent emissivity was further extended to real-time engine control by removing burst signals at real time (l MHz) and at engine DEC data rate. Simulink models are used to simulate the performances of the filter designs under engine normal and burst conditions. The results are compared with current LOS pyrometer results and show great advantage. The proposed model enables new LOS pyrometer design for improved engine control over wide range of operating conditions. [ABSTRACT FROM AUTHOR]

The ingestion of a vortex by an aero-engine is potentially an area of concern for current and future aircraft-engine configurations. However, there are very little experimental data on the characteristics of a streamwise vortex undergoing ingestion through a contracting streamtube. To address this dearth of knowledge, the ingestion of a streamwise vortex has been studied experimentally using stereoscopic particle image velocimetry (stereo PIV). A subscale model of an aircraft intake has been used to generate a contracting capture streamtube, and an isolated streamwise vortex has been generated upstream of the intake using semispan NACA 0012 and delta wing vortex generators (VGs). A range of contraction ratios, vortex Reynolds numbers, and vortex initial conditions have been examined. Measurements on planes perpendicular to the freestream flow show that the vortex undergoes notable levels of intensification through the contraction streamtube. The characteristics of the vortex are dependent on the streamtube contraction level, the initial aerodynamic characteristics of the vortex, and the trajectory that the vortex follows inside the capture streamtube. Results from inviscid, incompressible vortex filament theoty have been compared with the experimental data. At relatively low streamtube contraction ratios this theory provides a good estimate of the vortex characteristics. However, at higher contraction levels, there are notable levels of diffusion, which render the vortex less intense than that anticipated from vortex filament theory. [ABSTRACT FROM AUTHOR]

A numerical evaluation of the effects of volcanic ash ingestion in a turbofan engine was carried out, with particular regard to the prediction of the erosion damage to fan blades. The ash concentration level examined in the study was below the flight limit because the aim of this study is to investigate the damage due to long-term exposure to low concentration levels. The work aims to the implementation of a numerical methodology that takes into account the geometry change of the fan blades during the exposure to volcanic ash. A dimensional and morphological characterization of a real volcanic ash sample from the Mount Etna volcano has been performed to model the particle flow dynamics using a computational fluid dynamics (CFD) code. The fan petformance in terms of the total piessute increase was calculated for both the baseline and damaged geometries to quantify the petformance deterioration trend with respect to the particle exposure time. For the calculation of the eroded fan performance, two different numerical approaches were considered. In the first approach, the erosion rate (ER) was evaluated based on the initial blade geometry and was held constant. In the second approach, the ER was updated as the erosion of the blade continued. The second approach shows a higher deterioration of the pressure rise across the fan, suggesting that the variation of the ER due to the blade shape modification cannot be neglected in the calculations [ABSTRACT FROM AUTHOR]

The paper presents the use of different approaches to engine health assessment using onwing data obtained over a year from an engine of a commercial short-range aircraft. The on-wing measurements are analyzed with three different approaches, two of which employ two models of different quality. Initially, the measurements are used as the sole source of information and are postprocessed utilizing a simple "model" (a table of corrected parameter values at different engine power levels) to obtain diagnostic information. Next, suitable engine models are built utilizing a semi-automated method which allows for quick and efficient creation of engine models adapted to specific data. Two engine models are created, one based on publicly available data and one adapted to engine specific on-wing "healthy" data. These models of different details are used in a specific diagnostic process employing model-based diagnostic methods, namely the probabilistic neural network (PNN) method and the deterioration tracking method. The results demonstrate the level of diagnostic information that can be obtained for this set of data from each approach (raw data, generic engine model or adapted to measurements engine model). A subsystem fault is correctly identified utilizing the diagnostic process combined with the engine specific mode! while the deterioration tracking method provides additional information about engine deterioration. [ABSTRACT FROM AUTHOR]

Aircraft engines customarily implement squeeze film clampers (SFDs) to dissipate mechanical energy caused by rotor vibration and to isolate the rotor from its structural frame. The paper presents experimental results for the dynamic forced performance of an open ends SFD operating with large amplitude whirl motions, centered and off-centered. The test rig comprises of an elastically supported bearing with a damper section, 127 mm in diameter, having two parallel film lands separated by a central groove. Each film land is 25.4 mm long with radial clearance c = 0.251 mm. The central groove, 12.7 mm long, has a depth of 9.5 mm (38c). An ISO VG 2 lubricant flows into the groove via three 2.5 mm orifices, 120 deg apart, and then passes through the film lands to exit at ambient condition. Two orthogonally placed shakers apply dynamic loads on the bearing to induce circular orbit motions with whirl frequency ranging from 10 Hz to 100 Hz. A static loader, 45 deg away from each shaker, pulls the bearing to a static eccentricity (es). Measurements of dynamic loads and the ensuing bearing displacements and accelerations, as well as the film and groove dynamic pressures, were obtained for eight orbit amplitudes (r = 0.08c to ~0.71c) and under four static eccentricities (es = 0.0c to ~0.76c). The experimental damping coefficients increase quickly as the bearing offset increases (es/c → 0.76) while remaining impervious to the amplitude of whirl orbit (r/c → 0.51). The inertia coefficients decrease rapidly as the orbit amplitude grows large, r> 0.51c, but increase with the static eccentricity. A comparison with test results obtained with an identical damper hut having a smaller clearance (cs = 0.141 mm) (San Andres, L., 2012, "Damping and Inertia Coefficients for Two Open Ends Squeeze Film Dampers With a Central Groove: Measurements and Predictions," ASME J. Eng. Gas Turbines Power, 134(10), p. 102506), show the prior damping and inertia coefficients are larger, ~5.0 and ~2.2 times larger than the current ones. These magnitudes agree modestly with theoretical ratios for damping and inertia coefficients scaling as (c/cs)³=5.7 and (c/cs = 1.8, respectively. In spite of the large difference in depths between a groove and a film land, the magnitudes of dynamic pressures recorded at the groove are similar to those in the lands. That is, the groove profoundly affects the dynamic forced response of the test damper. A computational physics model replicates the experimental whirl motions and predicts force coefficients spanning the same range of whirl frequencies, orbit radii, and static eccentricities. The model predictions reproduce with great fidelity the experimental force coefficients. The good agreement relies on the specification of an effective groove depth derived from one experiment. [ABSTRACT FROM AUTHOR]

According to the Lefebvre's model and flame volume (FV) concept, an FV model about lean blow-out (LBO) was proposed by authors in early study. On the other hand, due to the model parameter (FV) contained in FV model is obtained based on the experimental data, FV model could only be used in LBO analysis instead of prediction. In view of this, a hybrid FV model is proposed that combines the FV model with numerical simulation in the present study. The model parameters contained in the FV model are all estimated from the simulated nonreacting flows. Comparing with the experimental data for 11 combustors, the maximum and average uncertainties of hybrid FV model are ±16% and ±10%. [ABSTRACT FROM AUTHOR]

Lean-bum swirl stabilized combustors represent the key technology to reduce NOx emissions in modern aircraft engines. The high amount of air admitted through a lean-burn injection system is characterized by very complex flow structures, such as recirculations, vortex breakdown, and processing vortex core, which may deeply interact in the near wall region of the combustor liner. This interaction and its effects on the local cooling performance make the design of the cooling systems veiy challenging, accounting for the design and commission of new test rigs for detailed analysis. The main purpose of the present work is the characterization of the flow field and the wall heat transfer due to the interaction of a swirling flow coming out from real geometry injectors and a slot cooling system which generates film cooling in the first part of the combustor liner. The experimental setup consists of a nonreactive three sector planar rig in an open loop wind tunnel; the rig, developed within the EU project Low Emissions Core-Engine Technologies (LEMCOTEC), includes three swirlers, whose scaled geometry reproduces the real geometry of an Avio Aero partially evaporated and rapid mixing (PERM) injector technology, and a simple cooling scheme made up of a slot injection, reproducing the exhaust dome cooling mass flow. Test were carried out imposing realistic combustor operating conditions, especially in terms of reduced mass flow rate and pressure drop across the swirlers. The flow field is investigated by means of particle image velocimetiy (PIV), while the measurement of the heat transfer coefficient is performed through thermochromic liquid crystals (TLCs) steady state technique. PIV results show the behavior of flow field generated by the injectors, their mutual interaction, and the impact of the swirled main flow on the stability of the slot film cooling. TLC measurements, reported in terms of detailed 2D heat transfer coefficient maps, highlight the impact of the swirled flow and slot film cooling on wall heat transfer. [ABSTRACT FROM AUTHOR]

There has been renewed interest in the contrarotating (CR) fan configuration in aviation and other applications where size and weight are important design factors. Contrarotation recovers swirl energy compared with the single-rotor design, but this advantage is not fully harnessed due to, perhaps, the issue of noise. This study explores passive noise reduction for a small, axial-flow, CR fan with perj'orated trailing-edge for the upstream rotor and perforated leading-edge for the downstream rotor. The fan is designed with simple velocity triangle analyses, which are checked by 3D flow computations. The aerodynamic consequence and the acoustic benefit of such perforated blading are investigated experimentally. The results show that there is a reduction of total pressure compared with the baseline CR fan at the same rotating speeds, but this is easily compensated for by slightly raising the rotating speeds. A reduction of 6-7dB in overall noise is achieved for the same aerodynamic output, although there is a moderate noise increase in the high frequency range of 12.5-15.0 kHz due to blade perforations. The effect of inter-rotor separation distance is also investigated for the baseline design. A clear critical distance exists below which the increased spacing shows clear acoustic benefits. [ABSTRACT FROM AUTHOR]

The multivariable robust tracking control problem for aeroengine is considered in this paper. On the basis of the aeroengine nonlinear affine uncertain dynamic model, and according to uncertainty and disturbance estimator (UDE) control approach, a novel aeroengine multivariable robust nonlinear tracking control method is presented in order to provide favorable tracking and disturbance rejection performance. After getting a generalization form of UDE-based aeroengine multivariable controller, a simplification form of control law is obtained when a specified form of low-pass filter is applied. Reference model of the aeroengine system should have satisfying dynamic, thus an optional reference model is provided. Simulation on a twin-shaft aeroengine with two inputs, verifies the effectiveness of the proposed method. [ABSTRACT FROM AUTHOR]

Aerodynamic instabilities such as stall and surge may occur in compressors, possibly leading to mechanical failures so their avoidance is crucial. A better understanding of those phenomena and an accurate prediction are necessary to improve both the performance and the safety. A surge event in a compressor threatens the mechanical integrity of the aircraft engine, and this remains true for a research compressor on a test rig. As a result, few experimental data on surge are available. Moreover, there are technological, restrictive constraints that exist on test rigs and limit severely the type of data obtainable experimentally. This partially explains why numerical simulation has become a usual, complementaiy and convenient tool to collect data in a compressor, as it does not disturb the flow nor does it encounter technological limits. Despite the inherent difficulties, an entire surge cycle has been simulated in a high-speed, high-pressure, multistage research compressor, using an implicit, time-accurate, 3D compressible unsteady Reynoldsaveraged Navier-Stokes solver. First, the paper presents the main features of the surge cycle obtained, along with those from the experimental cycle, for a validation purpose. Four phases compose the surge cycle: surge inception, the reversed-flow phase, the recovery phase, and the repressurization of the compressor flow. All of them are described, and focus is put on surge inception and the reversed-flow phase, as they induce greater risk for the mechanical integrity of the machine. [ABSTRACT FROM AUTHOR]

The next generation of aircraft will face more challenging demands in both electrical and thermal loads. The larger thermal loads reduce the propulsion system efficiency by demanding bleed air from the main engine compressor or imposing a shaft load on the high or low pressure shaft. The approach adopted to power the thermal management system influences the overall fuel burn of the aircraft for a given mission. To assess these demands and to explore conceptual designs for the electrical and thermal management system, a dynamic vehicle level tip-to-tail (T2T) model has been developed. The T2T model captures and quantifies the energy exchanges throughout the aircraft. The following subsystems of the aircraft are simulated in the T2T model: air vehicle system, propulsion system, adaptive power thermal management system, fuel thermal management system, electrical system, and actuator system. This paper presents trade studies evaluating the impact of various approaches in power take-off from the main engine and approaches in control strategy. The trade studies identify different control strategies resulting in significant fuel savings for a given mission profile. [ABSTRACT FROM AUTHOR]

Recent technology reviews have identified the need for objective assessments of aircraft engine health management (EHM) technologies. To help address this issue, a gas path diagnostic benchmark problem has been created and made publicly available. This software tool, referred to as the Propulsion Diagnostic Method Evaluation Strategy (Pro-DIMES), has been constructed based on feedback provided by the aircraft EHM community. It provides a standard benchmark problem enabling users to develop, evaluate, and compare diagnostic methods. This paper will present an overview of ProDiMES along with a description of four gas path diagnostic methods developed and applied to the problem. These methods, which include analytical and empirical diagnostic techniques, will be described and associated blind-test-case metric results will be presented and compared. Lessons learned along with recommendations for improving the public benchmarking processes will also be presented and discussed. [ABSTRACT FROM AUTHOR]

A stable gain scheduled controller for a gas turbine engine that drives a variable pitch propeller is developed and described. A stability proof is developed for gain scheduled closed-loop system using global linearization and linear matrix inequality (LMI) techniques. Using convex optimization tools, a single quadratic Lyapunov function is computed for multiple linearizations near equilibrium and nonequilibrium points of the nonlinear closed-loop system. This approach guarantees stability of the closed-loop gas turbine engine system. To verify the stability of the closed-loop system on-line, an optimization problem is proposed, which is solvable using convex optimization tools. Simulation results show that the developed gain scheduled controller is capable to regulate a turboshaft engine for large thrust commands in a stable fashion with proper tracking performance. [ABSTRACT FROM AUTHOR]

An optimal baseline turbofan cycle designed for a performance level expected to be available around year 2050 is established. Detailed performance data are given in take-off, top of climb, and cruise to support the analysis. The losses are analyzed, based on a combined use of the first and second law of thermodynamics, in order to establish a basis for a discussion on future radical engine concepts and to quantify loss levels of very high performance engines. In light of the performance of the future baseline engine, three radical cycles designed to reduce the observed major loss sources are introduced. The combined use of a first and second law analysis of an open rotor engine, an intercooled recuperated engine, and an engine working with a pulse detonation combustion core is presented. In the past, virtually no attention has been paid to the systematic quantification of the irreversibility rates of such radical concepts. Previous research on this topic has concentrated on the analysis of the turbojet and the turbofan engine. In the developed framework, the irreversibility rates are quantified through the calculation of the exergy destruction per unit time. A striking strength of the analysis is that it establishes a common currency for comparing losses originating from very different physical sources of irreversibility. This substantially reduces the complexity of analyzing and comparing losses in aero engines. In particular, the analysis sheds new light on how the intercooled recuperated engine establishes its performance benefits. [ABSTRACT FROM AUTHOR]

Linear dynamic finite element analysis can be considered very reliable today for the design of aircraft engine components. Unfortunately, when these individual components are built into assemblies, the level of confidence in the results is reduced since the joints in the real structure introduce nonlinearity that cannot be reproduced with a linear model. Certain types of nonlinear joints in an aircraft engine, such as underplatform dampers and blade roots, have been investigated in great detail in the past, and their design and impact on the dynamic response of the engine is now well understood. With this increased confidence in the nonlinear analysis, the focus of research now moves towards other joint types of the engine that must be included in an analysis to allow an accurate prediction of the engine behavior. One such joint is the bolted flange, which is present in many forms on an aircraft engine. Its main use is the connection of different casing components to provide the structural support and gas tightness to the engine. This flange type is known to have a strong influence on the dynamics of the engine carcase. A detailed understanding of the nonlinear mechanisms at the contact is required to generate reliable models and this has been achieved through a combination of an existing nonlinear analysis capability and an experimental technique to accurately measure the nonlinear damping behavior of the flange. Initial results showed that the model could reproduce the correct characteristics of flange behavior, but the quantitative comparison was poor. From further experimental and analytical investigations it was identified that the quality of the flange model is critically dependent on two aspects: the steady stress/load distribution across the joint and the number and distribution of nonlinear elements. An improved modeling approach was developed that led to a good correlation with the experimental results and a good understanding of the underlying nonlinear mechanisms at the flange interface. [ABSTRACT FROM AUTHOR]

In modern turbomachinery, abradable materials are implemented on casings to reduce operating tip clearances and mitigate direct unilateral contact occurrences between rotating and stationary components. However, both experimental and numerical investigations revealed that blade / abradable interactions may lead to blade failures. In order to comprehend the underlying mechanism, an accurate modeling of the abradable removal process is required. Time-marching strategies where the abradable removal is modeled through plasticity are available but another angle of attack is proposed in this work. It is assumed that the removal of abradable liners shares similarities with machine tool chatter encountered in manufacturing. Chatter is a self-excited vibration caused by the interaction between the machine and the workpiece through the cutting forces and the corresponding dynamics are efficiently captured by delay differential equations. These equations differ from ordinary differential equations in the sense that previous states of the system are involved in the formulation. This mathematical framework is employed here for the exploration of the blade stability during abradable removal. The proposed tool advantageously features a reduced computational cost and consistency with existing time-marching solution methods. Potentially dangerous interaction regimes are accurately predicted and instability lobes match both the flexural and torsional modal responses. Essentially, the regenerative nature of chatter in machining processes can also be attributed to abradable coating removal in turbomachinery. [ABSTRACT FROM AUTHOR]

Large-eddy simulation (LES) is applied to turbulent spray combustion fields in a subscale (1/2) aircraft jet engine combustor with an air-blast type swirl fuel nozzle and validity is examined by comparing with measurements. In the LES, Jet-A is used as liquid fuel, and individual droplet motion is tracked in a Lagrangian manner with a parcel model. As a turbulent combustion model, the extended flamelet/progress-variable approach, in which heat transfer between droplets and ambient gas including radiation and heat loss from walls can be taken into account, is employed. A detailed chemistry mechanism of Jet-A with 1537 reactions and 274 chemical species is used. The radiative heat transfer is computed by the discrete ordinate (DO) method. The equivalence ratio ranges from 0.91 to 1.29. The comparisons of the predicted droplet velocity and size, gaseous temperature, NO, and soot emissions with the measurements show that the present LES is capable of capturing the general features of the turbulent spray combustion fields in the subscale (1/2) aircraft jet engine combustor. [ABSTRACT FROM AUTHOR]

A Kalman filter-based approach for integrated on-line aircraft engine performance estimation and gas path fault diagnostics is presented. This technique is specifically designed for underdetermined estimation problems where there are more unknown system parameters representing deterioration and faults than available sensor measurements. A previously developed methodology is applied to optimally design a Kalman filter to estimate a vector of tuning parameters, appropriately sized to enable estimation. The estimated tuning parameters can then be transformed into a larger vector of health parameters representing system performance deterioration and fault effects. The results of this study show that basing fault isolation decisions solely on the estimated health parameter vector does not provide ideal results. Furthermore, expanding the number of the health parameters to address additional gas path faults causes a decrease in the estimation accuracy of those health parameters representative of turbomachinery performance deterioration. However, improved fault isolation performance is demonstrated through direct analysis of the estimated tuning parameters produced by the Kalman filter. This was found to provide equivalent or superior accuracy compared to the conventional fault isolation approach based on the analysis of sensed engine outputs, while simplifying online implementation requirements. Results from the application of these techniques to an aircraft engine simulation are presented and discussed. [ABSTRACT FROM AUTHOR]

Magnetic resonance imaging experiments have provided the three-dimensional mean concentration and three component mean velocity field for a typical trailing edge film-cooling cutback geometry built into a conventional uncambered airfoil. This geometry is typical of modern aircraft engines and includes three dimensional slot jets separated by tapered lands. Previous analysis of these data identified the critical mean flow structures that contribute to rapid mixing and low effectiveness in the fully turbulent flow. Three new trailing edge geometries were designed to modify the large scale mean flow struc-tures responsible for suiface effectiveness degradation. One modification called the Dol-phin Nose attempted to weaken strong vortex flows by reducing three dimensionality near the slot breakout. This design changed the flow structure but resulted in minimal improvement in the surface effectiveness. Two other designs called the Shield and Rounded Shield changed the land planform and added an overhanging land edge while maintaining the same breakout surface. These designs substantially modified the vortex structure and improved the suiface effectiveness by as much as 30%. Improvements included superior coolant uniformity on the breakout surface which reduces potential thermal stresses. The utilization of the time averaged data from combined magnetic reso-nance velocimetry (MRV) and concentration (MRC) experiments for designing improved trailing edge breakout film cooling is demonstrated. [ABSTRACT FROM AUTHOR]

Crackle noise from heated supersonic jets is characterized by the presence of strong positive pressure impulses resulting in a strongly skewed far-field pressure signal. These strong positive pressure impulses are associated with N-shaped waveforms involving a shock like compression and, thus, is very annoying to observers when it occurs. Unlike broadband shock-associated noise which dominates at upstream angles, crackle reaches a maximum at downstream angles associated with the peak jet noise directivity. Recent experiments (Martens et al., 2011, "The Effect of Chevrons on Crackle--Engine and Scale Model Results," Proceedings of the ASME Turbo Expo, Paper No. GT2011-46417) have shown that the addition of chevrons to the nozzle lip can significantly reduce crackle, especially in full-scale high-power tests. Because of these observations, it was conjectured that crackle is associated with coherent large scale flow structures produced by the baseline nozzle and that the formation of these structures are interrupted by the presence of the chevrons, which leads to noise reduction. In particular, shocklets attached to large eddies are postulated as a possible aerodynamic mechanism for the formation of crackle. In this paper, we test this hypothesis through a high-fidelity large-eddy simulation (LES) of a hot supersonic jet of Mach number 1.56 and a total temperature ratio of 3.65. We use the LES solver CHARLES developed by Cascade Technologies, Inc., to capture the turbulent jet plume on fully unstructured meshes. [ABSTRACT FROM AUTHOR]

A tangential sump design is employed in several aero-engines mainly due to its simplicity and compactness. However various problems have prompted investigations of the sump design. An experimental program at Purdue University (Chandra, 2006, "Flows in Turbine Engine Oil Sumps," PhD. thesis, School of Aeronautics and Astronautics, Purdue University and Radocaj, 2001, "Experimental Characterization of a Simple Gas Turbine Engine Sump Geometry," M.S. thesis, School of Aeronautics and Astronautics, Purdue University) was conducted to investigate the characteristics of a tangential sump design like one used in the AE3007 aero-engine. The research employed a transparent chamber to allow unprecedented view of the flow inside the bearing chamber. It was quickly found that a persistent liquid pooling near the drain entrance of a tangential sump obstructs the outflow to the scavenge pump. The airflow near the drain entrance has been shown to have strong reverse flow, even with, or perhaps due to, substantial over-scavenging. Various modifications were tested with varying degrees of success. The work highlights limitations of a tangential sump design. Based on the lessons learned, a new sump design was proposed. Preliminary test of the new sump design has shown positive results. [ABSTRACT FROM AUTHOR]

The Japan Aerospace Exploration Agency (JAXA) is conducting research and development on aircraft engine technologies to reduce environmental impact for the Technology Development Project for Clean Engines (TechCLEAN). As a part of the project, combustion technologies have been developed with an aggressive target that is an 80% reduction over the NOx threshold of the International Civil Aviation Organization (ICAO) Committee on Aviation Environmental Protection (CAEP)/4 standard. A staged fuel nozzle with a pilot mixer and a main mixer was developed and tested using a single-sector combustor under the target engine's landing and takeoff (LTO) cycle conditions with a rated output of 40kN and an overall pressure ratio of 25.8. The test results showed a 77% reduction over the CAEPI4 NOx standard. However, the reduction in smoke at thrust conditions higher than the 30% MTO condition and of CO emission at thrust conditions lower than the 85% MTO condition are necessary. In the present study, an additional fuel burner was designed and tested with the staged fuel nozzle in a single-sector combustor to control emissions. The test results show that the combustor enables an 82% reduction in NOx emissions relative to the ICAO CAEPI4 standard and a drastic reduction in smoke and CO emissions. [ABSTRACT FROM AUTHOR]

The finite element analysis of individual components of aircraft engine casings provides high accuracy and a good agreement with the measured response data. However, when these components are assembled, the accuracy of such predictions can significantly deteriorate since models describing stiffness and friction properties of joints are linearized. A full nonlinear analysis of the casing flanges is required to fully include the influence of the bolted joints, model the flexibility in the contact interface, and consider the nonlinear behavior of the contact due to partial slip and separation. In this paper different nonlinear models of casings are investigated with an available nonlinear analysis tool: A parametric study of the contact interface meshes is conducted to identify a satisfying analysis approach. The dynamic flange behavior is analyzed in detail, including effects of the bolt and normal load distribution. A comparison of the introduced nonlinear modeling with more traditional rigid or linear-elastic flange joint models is carried out to evaluate the effect of the nonlinear approach. The study demonstrates the nonlinear nature of a casing flange joint and highlights the need to include them in the analysis. The detailed modeling of the contact interaction of joints gives an insight in the nonlinear contact behavior of flanges of aircraft engine casings, and the predictive capabilities for the nonlinear analysis of gas turbine engines. [ABSTRACT FROM AUTHOR]

This paper presents a method of modeling contra-rotating turbomachinery components for engine performance simulations. The first step is to generate the performance characteristics of such components. In this study, suitably modified one-dimensional mean line codes are used. The characteristics are then converted to three-dimensional tables (maps). Compared to conventional turbomachinery component maps, the speed ratio between the two shafts is included as an additional map parameter and the torque ratio as an additional table. Dedicated component models are then developed that use these maps to simulate design and off-design operation at the component and engine levels. Using this approach, a performance model of a geared turbofan with a contra-rotating core (CRC) is created. This configuration was investigated in the context of the European program "NEW Aero-Engine Core Concepts" (NEWAC). The core consists of a seven-stage compressor and a two-stage turbine without interstage stators and with successive rotors running in the opposite direction through the introduction of a rotating outer spool. Such a configuration results in a reduced parts count, length, weight, and cost of the entire high pressure (HP) system. Additionally, the core efficiency is improved due to reduced cooling airflow requirements. The model is then coupled to an aircraft performance model and a typical mission is carried out. The results are compared against those of a similar configuration employing a conventional core and identical design point performance. For the given aircraft-mission combination and assuming a 10% engine weight saving when using the CRC arrangement over the conventional one, a total fuel burn reduction of 1.1% is predicted. [ABSTRACT FROM AUTHOR]

Aircraft engine rotors are particularly sensitive to rotor imbalance and sudden maneuver loads, since they are always supported on rolling element bearings with little damping. Most engines incorporate squeeze film dampers (SFDs) as means to dissipate mechanical Energy from rotor vibrations and to ensure system stability. The paper quantifies experimentally the forced performance of a SFD comprising two parallel film lands separated by a deep central groove. Tests are conducted on two open ends SFDs, both with diameter D = 127mm and nominal radial clearance c = 0.127mm. One damper has film lands with length L = 12.7ram (short length), while the other has 25.4mm land lengths. The central groove has width L and depth 3/4 L. A light viscosity lubricant flows into the central groove via three orifices, 120 deg apart and then through the film lands to finally exit to ambient. In operation, a static loader pulls the bearing to various eccentric positions and electromagnetic shakers excite the test system with periodic loads to generate whirl orbits of specific amplitudes. A frequency domain method identifies the SFD damping and inertia force coefficients. The long damper generates six times more damping and about three times more added mass than the short length damper. The damping coefficients are sensitive to the static eccentricity (up to ∼0.5 c), while showing lesser dependency on the amplitude of whirl motion (up to 0.2 c). On the other hand, inertia coefficients increase mildly with static eccentricity and decrease as the amplitude of whirl motion increases. Cross-coupled force coefficients are insignificant for all imposed operating conditions on either damper. Large dynamic pressures recorded in the central groove demonstrate the groove does not isolate the adjacent squeeze film lands, but contributes to the amplification of the film lands' reaction forces. Predictions from a novel SFD model that includes flow interactions in the central groove and feed orifices agree well with the test force coefficients for both dampers. The test data and predictions advance current knowledge and demonstrate that SFD-forced performance is tied to the lubricant feed arrangement. [ABSTRACT FROM AUTHOR]

Model-based design has attracted much attention in the field of aircraft engine control in recent years. As an aircraft engine is a complicated thermomechanical system, a nonlinear process model can only represent it. This necessitates the study of the nonlinear control techniques. Based on our recent results, this paper proposes a novel design approach based on a generalized Gronwall-Bellman lemma. Important results are obtained on, bounding behavior of the nonlinear states of the engine. The proposed method is easy to design and tune with the appealing feature of enlarging the feasible control envelope. Finally, a simulation study is provided to validate the effectiveness of the control design approach. [ABSTRACT FROM AUTHOR]

A direct method of aircraft engine health monitoring is introduced that can isolate a degraded engine component in-flight. The method utilizes continuous wavelet transformations of transient engine sensory data to represent their shape attributes in the time scale domain. This enables contrasting the shapes of the current engine outputs with those previously collected from the engine. Continuous wavelet transforms also provide enhanced delineation of the engine transients in the time scale domain. This enables identification of minute differences between the outputs affected by component degradations and between the sensitivities of modeled outputs with respect to the health parameters or components. The presence of these differences is used in this method as evidence of degradation effects on the outputs and/or parameters or components effects on the outputs. The effectiveness of the proposed method is evaluated in engine simulations. The results indicate that with the suite of outputs currently available on-board 70% to 96% of the degraded components simulated can be isolated for new and older engines. [ABSTRACT FROM AUTHOR]

Aircraft exhaust contains nonvolatile (soot) particulate matter (PM), trace gas pollutants, and volatile PM precursor material. Nonvolatile soot particles are predominantly present at the engine exit plane, but volatile PM precursors form new particles or add mass to the existing ones as the exhaust is diluted and cooled. Accurately characterizing the volatile PM mass, number, and size distribution is challenging due to this evolving nature and the impact of local ambient conditions on the gas-to-particle conversion processes. To accurately and consistently measure the aircraft PM emissions, a dilution and aging sampling system that can condense volatile precursors to particle phase to simulate the atmospheric evolution of aircraft engine exhaust has been developed. In this paper, a field demonstration of its operation is described. The dilution/aging probe system was tested using both a combustor rig and on-wing CFM56-7 engines. During the combustor rig testing at NASA Glenn Research Center, the dilution/aging probe supported formation of both nucleation/growth mode particles and soot coatings. The results showed that by increasing residence time, the nucleation particles become larger in size, increase in total mass, and decrease in number. During the on-wing CFM56-7 engine testing at Chicago Midway Airport, the dilution/aging probe was able to form soot coatings along with nucleation mode particles, unlike conventional 1-m probe engine measurements. The number concentration of nucleation particles depended on the sample fraction and relative humidity of the dilution air. The performance of the instrument is analyzed and explained using computational microphysics simulations. [ABSTRACT FROM AUTHOR]