A Fully Coupled Approach for the Integration of 3D-CFD Component Simulation in Overall Engine Performance Analysis

Due to a high degree of complexity and computational effort, overall system simulations of jet engines are typically performed as 0-dimensional thermodynamic performance analysis. Within these simulations and especially in the early cycle design phase, the usage of generic component characteristics is common practice. Of course these characteristics often cannot account for true engine component geometries and operating characteristics which may cause serious deviations between simulated and actual component and overall system performance. This leads to the approach of multi-fidelity simulation, often referred to as zooming, where single components of the thermodynamic cycle model are replaced by higher-order procedures. Hereby the consideration of actual component geometries and performance in an overall system context is enabled and global optimization goals may be considered in the engine design process. The purpose of this study is to present a fully automated approach for the integration of a 3D-CFD component simulation into a thermodynamic overall system simulation. As a use case, a 0D-performance model of the IAE-V2527 engine is combined with a CFD model of the appropriate fan component. The methodology is based on the DLR in-house performance synthesis and preliminary design environment GTlab combined with the DLR in-house CFD solver TRACE. Both, the performance calculation as well as the CFD simulation are part of a fully automated process chain within the GTlab environment. The exchange of boundary conditions between the different fidelity levels is accomplished by operating both simulation procedures on a central data model which is one of the essential parts of GTlab. Furthermore iteration management, progress monitoring as well as error handling are part of the GTlab process control environment. Based on the CFD results comprising fan efficiency, pressure ratio and mass flow, a map scaling methodology as it is commonly used for engine condition monitoring purposes is applied within the performance simulation. Hereby the operating behavior of the CFD fan model can be easily transferred into the overall system simulation which consequently leads to a divergent operating characteristic of the fan module. For this reason, all other engine components will see a shift in their operating conditions even in case of otherwise constant boundary conditions. The described simulation procedure is carried out for characteristic operating conditions of the engine.

Development and Testing of a Gas Turbine Engine Combustion Air Inlet Filtration System for the USMC Amphibious Combat Vehicle

The Auxiliary Ships and New Acquisition Support Branch (Code 425) of the Naval Surface Warfare Center, Philadelphia Division conducted a study to assist the Marine Corps Systems Command in assessing the feasibility of using a gas turbine engine as a propulsion system on future United States Marine Corps Amphibious Combat Vehicles (ACV). The study was focused on developing and testing a gas turbine intake solution for the ACV that can remove saltwater from the intake airstream of a notional 3,000 horsepower ACV engine. Code 425 developed a two-part solution for the intake of the ACV. The first part of the solution is an intake shroud designed to elevate the intake to protect the engine from deck water wash. The second part of the solution is the Combustion Air Separation System (CASS), a gas turbine intake filtration system designed to remove marine contaminants that enter the intake. Code 425 tested a CASS prototype for its efficiency at removing saltwater spray and bulk water up to 10 gallons per minute. Test results showed that the CASS met each requirement and that an ACV intake system incorporating both the intake shroud and the CASS should protect the gas turbine engine from saltwater ingestion.

A Semantic Differential for Evaluating the Sound Quality of Fan Systems

Fans are main components e.g. in heating, ventilating and air conditioning systems for vehicles or buildings, cooling units of engines and electronic circuits, and household appliances such as kitchen exhaust hoods or vacuum cleaners. End-users increasingly demand a high sound quality of their system or device. The overall objective of a recent research project at the University of Siegen is a multidimensional assessment of fan sound quality. In a first step an advanced novel semantic differential for the assessment of fan-related sounds is established with the aid of carefully designed jury tests. Eventually, this semantic differential is employed for sound quality jury tests of fans in kitchen exhaust hoods, heat pumps and air purifiers as a first case. Finally, a prediction model is suggested, which relates the outcome from the jury tests to objective metrics. A principal component analysis is carried out and yields five main assessment criteria with 23 relevant adjective scales. The results show that the perceived sound quality of fan systems is mainly determined by the loudness and tonality of the sound. The spectral content (represented by the sharpness) as well as the time structure (represented by the roughness) have no significant impact on perceived sound quality of the fan systems investigated.

Analysis of a Large-Scale Cooling System Fan Gearbox Loads

South Africa’s coal-fired power stations use super heated steam to drive generator turbines. In arid regions, air-cooled condensers (ACCs) are used to condense the process steam. These ACCs consists of an array of over 200 axial flow fans, each driven by a motor via a reduction gearbox. Distorted fan inlet air flow conditions cause transient blade loading, which results in variations in output shaft bending and torque. A measurement project was conducted where the input and output shaft of such a gearbox were instrumented with strain gauges and wireless bridge amplifiers. Gearbox shaft speed and vibration were also measured. Torsional and bending strains were measured for a variety of operational conditions, where correlations were seen between gearbox loading and wind conditions. The input side experienced no unexpected loads from the motor or changing wind conditions, whereas output shaft loading was influenced by the latter. Digital filters were applied to identify specific bending components, such as the influence of fan hub misalignment and dynamic blade loading. Reverse loading of the gearbox was measured during the fan stop period, and vibration analysis revealed torsional and gearbox vibrations. This investigation documented reliable full scale ACC gearbox loads.

The Design of a Large Diameter Axial Flow Fan for Air-Cooled Heat Exchanger Applications

An axial flow fan design methodology is developed to design large diameter, low pressure rise, rotor-only fans for large air-cooled heat exchangers. The procedure aims to design highly efficient axial flow fans that perform well when subjected to off design conditions commonly encountered in air-cooled heat exchangers. The procedure makes use of several optimisation steps in order to achieve this. These steps include optimising the hub-tip ratio, vortex distribution, blading and aerofoil camber distributions in order to attain maximum total-to-static efficiency at the design point. In order to validate the design procedure a 24 ft, 8 bladed axial flow fan is designed to the specifications required for an air-cooled heat exchanger for a concentrated solar power (CSP) plant. The designed fan is numerically evaluated using both a modified version of the actuator disk model and a three dimensional periodic fan blade model. The results of these CFD simulations are used to evaluate the design procedure by comparing the fan performance characteristic data to the design specification and values calculated by the design code. The flow field directly down stream of the fan is also analysed in order to evaluate how closely the numerically predicted flow field matches the designed flow field, as well as determine whether the assumptions made in the design procedure are reasonable. The fan is found to meet the required pressure rise, however the fan total-to-static efficiency is found to be lower than estimated during the design process. The actuator disk model is found to under estimate the power consumption of the fan, however the actuator disk model does provide a reasonable estimate of the exit flow conditions as well as the total-to-static pressure characteristic of the fan.

Preliminary Investigation on the Effect of the Modification of the Sweep Angle at the Blade Tip of Forward Swept Axial Fans

The flow path close to the suction side of fan rotor blades mostly affects the overall drag of the blading. The blade lift is affected as well because of the separation of the low energy boundary layer that drives the blade into stall at low fan flow rates. Forward sweep allows to position the airfoil sections of blades featuring a positive circulation gradient along the span so that they “accompany” the near-wall flow trajectories at the blade suction side. So, rotor efficiency and stall margin of the fan can be improved. On the other hand, blade end effects play a relevant role in high hub-to-tip and low aspect ratio rotors and may compromise the effectiveness of forward sweep. Nevertheless, some authors in the literature stated the beneficial contribution of changing the sweep angle at the ends of the blade both at design and off-design conditions. The paper studies the end effects on constant-swirl design rotors by means of CFD simulations focusing on the distribution of blade sweep in the near-tip region. In particular, the performance and efficiency calculated for a forward swept tube-axial fan featuring a hub-to-tip ratio equal to 0.4 are compared with those estimated for the corresponding unswept fan at equal duty point. Several modifications of the sweep distribution in the blade tip region are considered in the swept fan to quantify their effect on performance, efficiency and stall margin. Results show that the addition of up to 6 degrees of local forward sweep at the blade tip to the unswept blading does not affect fan pressure at design operation. On the other hand, this local increase of the sweep angle allows for a very notable increase of the peak pressure and efficiency at flow rates close to stall inception.

A 3D Shape Design and Optimization Method for Natural Laminar Flow Nacelle

With the rapid development of high bypass ratio turbofan engine, the proportion of the nacelle drag increases obviously in the total drag of the aircraft with the increase of nacelle surface area. And the frictional resistance is one of the major contributors of drag. Under the same Reynolds number, the friction resistance in turbulent boundary layer is about 10 times larger as that in laminar boundary layer. Therefore, a correctly profiled engine nacelle will delay the transition in the boundary layer and allow laminar flow to extend back, resulting in a substantial drag reduction. In the previous conference paper (9th reference), a 2D nacelle longitudinal profile-line geometry generator, which allows curvature and slope-of-curvature to be continuous was developed and presented. This established an optimization system to minimize nacelle frictional drag. One of the nacelle profile-line is optimized to achieve minimum drag coefficient, and then is stacked with the other original profile-lines to form the 3D isolated nacelle aerodynamic shape. Finally, a total 23% of nacelle outer surface maintains a laminar flow and its frictional drag coefficient is less than initial shape. This paper proposes a new 2D nacelle longitudinal profile-line design method, based on PARSEC parameterization, with can generate the profile-line rapidly and precisely. Conservation Full Potential Equation was used to calculate the aerodynamic distribution and obtain the transition location. Then adaptive simulated annealing genetic algorithm was adapted to search 2D profiles of low drag, which would be applied to narrow down design space in 3D nacelle optimization. Second, 2D profiles were stacked circumferentially, by NURBS surface generator, to form the 3D nacelle aerodynamic shape, and an optimization system was established, in combination with the 3D nacelle generator, γ–Reθ transition model, Kriging surrogate model and adaptive simulated annealing algorithm, for natural laminar flow nacelle design. Finally, a total 34% of nacelle surface maintains a laminar flow and its frictional drag coefficient is less than initial shape. The generated optimized loft was evaluated by CFD to determine if the low drag of this optimized nacelle shape can be maintained under different Mach numbers and angles of attack.

Engine Fleet-Management: The Use of Digital Twins From a MRO Perspective

Engine operating cost is a major contributor to the direct operating cost of aircraft. Therefore, the minimization of engine operating cost per flight-hour is a key aspect for airlines to operate successfully under challenging market conditions. The interaction between maintenance cost, operating cost, asset value, lease and replacement cost describes the area of conflict in which engine fleets can be optimized. State-of-the-art fleet management is based on advanced diagnostic and prognostic methods on engine and component level to provide optimized long-term removal and work-scoping forecasts on fleet level based on the individual operation. The key element of these methods is a digital twin of the active engines consisting of multilevel models of the engine and its components. This digital twin can be used to support deterioration and failure analysis, predict life consumption of critical parts and relate the specific operation of a customer to the real and expected condition of the engines on-wing and at induction to the shop. The fleet management data is constantly updated based on operational data sent from the engines as well as line maintenance and shop data. The approach is illustrated along the real application on the CFM56-5C, a mature commercial two-spool high bypass engine installed on the Airbus A340-300. It can be shown, that the new methodology results in major improvements on the considered fleets.

Flow Interactions Between Shrouded Power Turbine and Nonaxisymmetric Exhaust Volute for Marine Gas Turbines

The marine gas turbine exhaust volute is an important component that connects a power turbine and an exhaust system, and it is of great importance to the overall performance of the gas turbine. Gases exhausted from the power turbine are expanded and deflected 90 degrees in the exhaust volute, and then discharge radially into the exhaust system. The flows in the power turbine and the nonaxisymmetric exhaust volute are closely coupled and inherently unsteady. The flow interactions between the power turbine and the exhaust volute have a significant influence on the shrouded rotor blade aerodynamic forces. However, the interactions have not been taken into account properly in current power turbine design approaches. The present study aims to investigate the flow interactions between the last stage of a shrouded power turbine and the nonaxisymmetric exhaust volute with struts. Special attention is given to the coupled aerodynamics and pressure response studies. This work was carried out using coupled computational fluid dynamics (CFD) simulations with the computational domain including a stator vane, 76 shrouded rotor blades, 9 struts and an exhaust volute. Three-dimensional (3D) unsteady and steady Reynolds-averaged Navier-Stokes (RANS) solutions in conjunction with a Spalart-Allmaras turbulence model are utilized to investigate the aerodynamic characteristics of shrouded rotors and an exhaust volute using a commercial CFD software ANSYS Fluent 14.0. The asymmetric flow fields are analyzed in detail; as are the unsteady pressures on the shrouded rotor blade. In addition, the unsteady total pressures at the volute outlet is also analyzed without consideration of the upstream turbine effects. Results show that the flows in the nonaxisymmetric exhaust volute are inherently unsteady; for the studied turbine-exhaust configuration the nonaxisymmetric back-pressure induced by the downstream volute leads to the local flow varying for each shrouded blade and low frequency fluctuations in the blade force. Detailed results from this investigation are presented and discussed in this paper.

Numerical Investigation of Effect of Inlet Distortion on Compressor Flow Field and Stability

The compressor aerodynamic design is conducted under the condition of clean inlet in general, but a compressor often operates under the condition of inlet distortion in the practical application. It has been proven by a lot of experimental and numerical investigations that inlet distortion can decrease the performance and stability of compressors. The circumferential or radial distorted inlet in mostly numerical investigations is made by changing the total pressure and total temperature in the inlet ring surface of the compressors. In most of inlet distortion experiments, distorted inlets are usually created by using wire net, flashboards, barriers or the generator of rotating distortion. The fashion of generating distorted inlet for experiment is different from that for numerical simulation. Consequently, the flow mechanism of affecting the flow field and stability of a compressor with distorted inlet for experiment is partly different than that for numerical simulation. In the numerical work reported here, the inlet distortion is generated by setting some barriers in the inlet ring surface of an axial subsonic compressor rotor. Two kinds of distorted inlet are investigated to exploring the effect of distorted range on the flow field and stability of the compressor with ten-passage unsteady numerical method. The numerical results show that the inlet distortions not only degrade the total pressure and efficiency of the compressor rotor, but also decrease the stability of the rotor. The larger the range of distorted inlet is, the stronger the adverse effect is. The comprehensive stall margin for the inlet distortion of 24 degrees and 48 degrees of ten-passages is reduced about 3.35% and 5.88% respectively. The detailed analysis of the flow field in the compressor indicates that the blockage resulted from tip clearance leakage vortex (TLV) and the flow separation near the suction surfaces of some blades tip for distorted inlet is more serious than that resulted from TLV for clean inlet. Moreover, the larger the range of distorted inlet is, the larger the range of the blockage is. The analysis of unsteady flow shows that during this process, which is that one rotor blade passes through the region affected by the distorted inlet, the range of the blockage in the rotor passage increases first, then reduces, and increases last.