ASME Conference Presenter Attendance Policy and Archival Proceedings

2016;():V001T00A001. doi:10.1115/GT2016-NS1.

This online compilation of papers from the ASME Turbo Expo 2016: Turbomachinery Technical Conference and Exposition (GT2016) represents the archival version of the Conference Proceedings. According to ASME’s conference presenter attendance policy, if a paper is not presented at the Conference by an author of the paper, the paper will not be published in the official archival Proceedings, which are registered with the Library of Congress and are submitted for abstracting and indexing. The paper also will not be published in The ASME Digital Collection and may not be cited as a published paper.

Commentary by Dr. Valentin Fuster

Aircraft Engine

2016;():V001T01A001. doi:10.1115/GT2016-56052.

The work described in this paper utilizes dust ingestion experimental results obtained using three Pratt/Whitney F-100, two GE F-101, one Pratt/Whitney J-57, and three Pratt/Whitney TF-33 military engines and two different combustor rigs (one utilizing a sector of the Pratt/Whitney F-100 annular combustor and the other utilizing an Allison T-56 can combustor) to scale results so that these previous experiments can be used to approximate the response of more current aircraft engines to foreign particle ingestion. Modern engines experience a combination of compression system erosion and material deposition in the combustor and on the high-pressure turbine inlet vanes (and rotor blade complications) whereas the older engines (P/W TF-33 and J-57) experienced primarily an erosion problem as a result of the lower turbine inlet temperatures. As part of the results presented in this paper the scaled estimates of material accumulation and component degradation have been compared to documented in-flight ash encounters, specifically KLM Flight 867, British Airways Flight 009, Qantas Flight 370, and a NASA scientific research flight. The results of the study allow one to make estimates of the time to initial issues for the RR RB-211, the GE CF-6, the GE/Snecma CFM-56, and the P/W JT9-D engines encountering dust clouds of specific concentration. Current engine certification procedures do not require any specific test condition that would approach the engine issues described in this paper.

Commentary by Dr. Valentin Fuster
2016;():V001T01A002. doi:10.1115/GT2016-56055.

The maximization of rapid fuel-air mixing is one of the essential issues for the efficient operation of scramjet engines. A delta wing with its height being 6mm is located ahead of the injector to enhance the mixing process between the injectant and air in the supersonic flow with the freestream Mach number being 3.75, and the influence of the distance between the delta wing and the injector on the mixing efficiency is evaluated numerically, as well as the effect of the jet-to-crossflow pressure ratio. At the same time, the predicted results obtained in the three-dimensional transverse injection flow field are compared with the available experimental data in the open literature, and the grid independency analysis is conducted as well. The obtained results show that the mixing efficiency increases with the decrease of the jet-to-crossflow pressure ratio, and this conclusion is consistent with that obtained in the transverse injection flow field. The predicted wall static pressure distributions show reasonable agreement with the experimental data, and the grid scale has only a slight impact on the predicted results. Further, it is observed that the mixing efficiency increase with the decrease of the distance between the delta wing and the injector, and the hydrogen penetrates deeper into the core flow when the distance is smaller. Accordingly, the plume area is larger. This illustrates that the transverse jet flow field is affected by the vortex generated by the delta wing, and the mixing process is enhanced. The maximum mixing efficiency at x = −350mm is nearly 0.84 in the range considered in this paper.

Commentary by Dr. Valentin Fuster
2016;():V001T01A003. doi:10.1115/GT2016-56059.

Particle ingestion into modern gas turbine engines is known to reduce performance and may damage many primary gas path components through erosion or deposition mechanisms. Many studies have been conducted that evaluate the effects of particulate ingestion in primary and secondary gas path components. However, modern gas turbines have gas path temperatures that are above most previous studies. As a result, this study performed particle deposition experiments at the Virginia Tech Aerothermal Rig facility at engine representative temperatures. Arizona Test Dust of 20 to 40 μm was chosen to represent the particle ingested into rotorcraft turbine engines in desert and sandy environments. The experimental setup impinged air and sand particles on a flat Hastelloy X coupon. The gas and sand mixture impacted the coupon at varying angles measured between the gas flow direction and coupon face, hereby referred to as coupon angle. For this study, gas and sand particles maintained a constant flow velocity of about 70 m/s and a temperature of about 1100°C. The coupon angle was varied between 30° to 90° for all experiments. The experimental results indicate sand deposition increased linearly from about 975 °C to 1075 °C for all coupon angles. A multiple linear regression model is used to estimate the amount of deposition that will occur on the test coupon as a function of gas path temperature and coupon angle. The model is adequate in explaining about 67% of the deposition that occurs for the tests. The remaining percentage could be explained with other factors such as particle injection rates and exact surface temperature where the deposits occur.

Commentary by Dr. Valentin Fuster
2016;():V001T01A004. doi:10.1115/GT2016-56062.

This work is the continuation of previous studies on gerotor-type pump performance in turbofan engine oil systems operated as feed pumps in single-phase liquid oil. The focus here is on scavenge pumps whose role is to pump a mix of air and oil. This paper is intended to present the modifications that had to be made on the test rig from the previous studies to model a scavenge system and more generally to add two-phase flow capacity. The paper presents results from the first successful experimental test campaign. The aim is to characterize the performance of a typical pump, already tested as a feed pump, in the scavenge system. The critical performance parameter studied is the volumetric efficiency which determines the size and weight of the pump. This paper ends by drawing conclusions on the rig and the results, and linking them with the previous single-phase flows studies.

Commentary by Dr. Valentin Fuster
2016;():V001T01A005. doi:10.1115/GT2016-56083.

A concept of mixing enhancement forced by rotating ramps has been presented, which based on the design of stationary ramps and Rim-Rotor Rotary Ramjet Engine. The commercial software Fluent has been conducted to study the nonreacting flowfields in a transverse hydrogen fuel injection channel. The results indicate a slowly increased mixing efficiency when the ramps speed increase. As the ramps rotate at a speed of 6000rpm, the mixing efficency is almost the same as the stationary ramps case. As the ramps rotate at a speed of 54295rpm, a 10%∼20% higher mixing efficiency is obtained.

Commentary by Dr. Valentin Fuster
2016;():V001T01A006. doi:10.1115/GT2016-56107.

Aero-engines incorporate various bearing chambers and these typically contain bearings, seals, rotating shafts, stationary walls and struts, and sometimes gears. Oil is supplied for lubrication and cooling and is removed (scavenged) from the sump region of the chamber (note that in some parts of the world the entire bearing chamber is referred to as the sump). Depending on the location and function of the bearing chamber, the sump region may be deep or shallow. Effective oil removal is essential as unnecessary working of the oil can lead to excessive heat generation and reduced overall efficiency. Therefore the design of the scavenge region in a bearing chamber, as well as the ability to assess its performance is very important.

Previous work, much of which was conducted at the University of Nottingham Technology Centre in Gas Turbine Transmission Systems (UTC) suggests that oil often does not flow cleanly into the off-take due to a combination of several factors: oil momentum, windage, three-dimensional air flow that blocks the off-take flow or transports oil away from the off-take, and pooling because of separated air flow that acts on the oil once oil momentum is dissipated.

Experimental research at the UTC found that scavenge performance is highly affected by the sump geometry, especially its depth. Variations of shallow sumps, although some are better than the others, cannot offer the same level of performance as a deep one. However space limitation in an engine often only allows for a shallow sump. This paper presents some experimental exploration on new design ideas. They are in the form of various inserts and attachments that were designed to improve scavenge performance of a shallow sump. These “custom” sumps were tested on the UTC’s scavenge test facility at various flow settings (wall film/flying droplets, liquid flow rates, scavenge ratios, shaft speeds). The residence volumes were measured and compared to a baseline configuration with reduction in residence volume desirable.

The inserts tested were a Grille Cover, a Stepped Spillway, a Perforated Plate and a Porous Insert. Both the Porous Insert and the Perforated Plate showed reduced residence volumes in the demanding droplet/windage dominated flow condition with the Perforated Plate offering the best improvement over baseline.

Commentary by Dr. Valentin Fuster
2016;():V001T01A007. doi:10.1115/GT2016-56174.

Fixed wing Micro Air Vehicles (MAVs) are widely powered by miniature brushless DC motors and mini propellers. These motors have low efficiencies in the MAVs’ operating range; hence consume more power and penalizing the endurance of MAVs flight. Mini propellers also suffer with lower efficiencies due to low Reynolds number effect compared to bigger propellers. Based on the power requirement of MAV’s; generally two bladed different size propellers are used to power these MAVs. Effect of Propeller slipstream wash on the lift generating wings is significant and it is well reported for bigger propellers. The strength of slipstream wash depends on the number of blades, diameter of the propeller, rotational speed, flight speed and trajectory. The effect of slipstream wash could be lowered by increasing the number of blades and with smaller diameter propellers. CSIR-NAL has designed, developed, and fabricated, efficient, 6inch diameter, two and three blade, light-weight, mini propellers using latest state of the art technological advancements for CSIR-NAL fixed wing MAV code named as Black Kite. Apart from this, these propellers are assessed for its realistic propulsive efficiencies using CSIR-NAL configured sophisticated precision test bench manufactured and supplied by M/s MAGTROL, Switzerland. The specialty of this test bench is that it can measure thrust, propellers shaft torque, input power and rotational speeds simultaneously. In the present study CSIR-NAL developed 6 inch diameter two and three bladed propellers of identical plan form are compared for their performance. The three bladed propellers generate 30% higher thrust by marginal weight and efficiency penalty, whereas the noise levels are reduced.

Topics: Blades , Propellers
Commentary by Dr. Valentin Fuster
2016;():V001T01A008. doi:10.1115/GT2016-56203.

The main role of the intake is to provide a sufficient mass flow to the engine face and sufficient flow homogeneity to the fan. Intake-fan interaction off design represents a critical issue in the design process because intake lines are set very early during the aircraft optimization.

The off-design operation of an aero-engine, strictly related to the intake flow field, can be mainly related to two different conditions. When the plane is in near ground position, vorticity can be ingested by the fan due to crosswind incidence. During the flight, distortions occur due to incidence. In these conditions, the windward lip is subjected to high acceleration followed by strong adverse pressure gradients, high streamline curvature and cohabitation of incompressible and transonic flow around the lip. All these features increase the risk of lip stall in flight at incidence or in crosswind near ground operation and increase the level of forcing seen by the fan blades because of the interaction with non-uniform flow from the intake.

This work deals with the study of two sources of distortions: ground vortex ingestion and flight at high incidence conditions. A test case representative of a current installation clearance from the ground has been investigated and the experimental data available in open literature validated the CFD approach. An intake, representative of a realistic civil aeroengine configuration flying at high incidence, has been investigated in powered and aspirated configurations. Distortion distributions have been characterized in terms of total loss distributions in space and in time. The beneficial effect of the presence of the fan in terms of distortion control has been demonstrated. The mutual effect between fan and incoming distortion from the intake has been assessed in terms of modal force and distortion control. CFD has been validated by means of comparisons between numerical results and experimental data have been provided. Waves predicted by CFD have been compared with an actuator disk approach prediction. The linear behavior of the lower disturbance frequency coming from distortion and the waves reflected by the fan has been demonstrated.

Commentary by Dr. Valentin Fuster
2016;():V001T01A009. doi:10.1115/GT2016-56227.

The design process of a gas turbine engine involves interrelated multi-disciplinary and multi-fidelity designs of engine components. Traditional component-based design process is not always able to capture the complicated physical phenomenon caused by component interactions. It is likely that such interactions are not resolved until hardware is built and tests are conducted. Component interactions can be captured by assembling all these components into one computational model. Nowadays, numerical solvers are fairly easy to use and the most time-consuming (in terms of man-hours) step for large scale gas turbine simulations is the preprocessing process. In this paper, a method is proposed to reduce its time-cost and make large scale gas turbine numerical simulations affordable in the design process. The method is based on a novel featured-based in-house geometry database. It allows the meshing modules to not only extract geometrical shapes of a computational model and additional attributes attached to the geometrical shapes as well, such as rotational frames, boundary types, materials, etc. This will considerably reduce the time-cost in setting up the boundary conditions for the models in a correct and consistent manner. Furthermore, since all the geometrical modules access to the same geometrical database, geometrical consistency is satisfied implicitly. This will remove the time-consuming process of checking possible mismatching in geometrical models when many components are present. The capability of the proposed method is demonstrated by meshing the whole gas path of a modern three-shaft engine and the Reynold’s Averaged Navier-Stokes (RANS) simulation of the whole gas path.

Commentary by Dr. Valentin Fuster
2016;():V001T01A010. doi:10.1115/GT2016-56297.

This paper presents experimental and numerical investigations of oil leakage across a conventional labyrinth seal commonly found in aero engine bearing chambers. Measurements and simulations were carried out in order to investigate the influence of chamber geometry and operating conditions on the reliability of the oil seal against leakage.

The main goal of the experiments was to determine a minimum required pressure difference Δpleak to prevent oil from leaving the bearing chamber for any given operating point. To determine this variable, the pressure inside the test rig was continuously lowered from a high pressure difference until oil was found to leave the bearing chamber. Using two pressure supplies, this pressure could be negative or positive. The results show that the minimum pressure depends on component design and rotational speed. While certain component designs may increase this pressure at low rotational speeds, thereby creating a safety margin for oil leakage, the opposite effect can manifest itself at higher rotational speeds.

Selected operating points were simulated using computational fluid dynamics employing the Volume-of-Fluid (VoF) approach. A comparison of the experimental and numerical results shows good qualitative agreement of the two phase flow phenomena inside the bearing chamber.

Commentary by Dr. Valentin Fuster
2016;():V001T01A011. doi:10.1115/GT2016-56313.

The Joule-/Brayton thermodynamic cycle is the base cycle of all major contemporary aero engines. Over the decades, the achievement of further significant improvements has become progressively challenging, and the increase of efficiency approaches physical limitations. In order to meet the ambitious long-term emission reduction targets, the introduction of radical new propulsion system concepts is indispensable. Various cycles promising significant efficiency improvements over the conventional Joule-/Brayton-cycle are being examined by the engine community. However, as no clear favorite has emerged from these potential technical solutions, a transparent methodological approach for the consistent evaluation of the concepts is necessary.

Consistent thermodynamic description and performance metrics for three engine cycles are presented in this paper: The turbofan as reference and two radical engine cycles, namely the composite cycle and the cycle-integrated parallel hybrid. Laws for the estimation of component performance for large parametric variations are introduced. A method for the estimation of power plant system mass for the investigated engine cycles is proposed to evaluate fuel burn reduction. The studies substantiated that the turbofan improvement potential is saturating. The composite cycle engine offers a tremendous potential for fuel burn improvement of 24.5% over state of the art turbofan engines, which allows meeting the emission reduction targets in 2035. The cycle-integrated parallel hybrid engine improves the turbofan moderately with year 2035 technology, but is not capable of meeting the corresponding emission reduction targets on a short-to-medium range aircraft platform.

Commentary by Dr. Valentin Fuster
2016;():V001T01A012. doi:10.1115/GT2016-56392.

The effects of the aft rotor on the inter-rotor flow field of an open rotor propulsion rig were examined. A Particle Image Velocimetry (PIV) dataset that was acquired phase locked to the front rotor position has been phase averaged based on the relative phase angle between the forward and aft rotors. The aft rotor phase was determined by feature tracking in raw PIV images through an image processing algorithm. The effects of the aft rotor potential field on the inter-rotor flow were analyzed and shown to be in reasonably good agreement with Computational Fluid Dynamics (CFD) simulations. The aft rotor position was shown to have a significant upstream effect, with implications for front rotor interaction noise. It was found that the aft rotor had no substantial effect on the position of the forward rotor tip vortex but did have a small effect on the circulation strength of the vortex when the rotors were highly loaded.

Commentary by Dr. Valentin Fuster
2016;():V001T01A013. doi:10.1115/GT2016-56540.

Inspired by Prandtl’s theory on aircraft wings with minimum induced drag, the authors introduced a double-bladed propeller, the Boxprop, intended for high-speed flight. The basic idea is to join the propeller blades pair-wise at the tip to improve aerodynamics and mechanical properties compared to the conventional propeller. The rather complex geometry of the double blades gives rise to new questions, particularly regarding the aerodynamics.

This paper presents a propeller wake energy analysis method which gives a better understanding of the potential performance benefits of the Boxprop and a means to improve its design.

CFD analysis of a five bladed Boxprop demonstrated its ability to generate typical levels of cruise thrust at a flight speed of Mach 0.75. The present work shows that the near tip velocity variations in the wake are weaker for this propeller than a conventional one, which is an indication that a counter rotating propeller designed with a Boxprop employed at the front may exhibit lower interaction noise.

Topics: Wakes , Propellers
Commentary by Dr. Valentin Fuster
2016;():V001T01A014. doi:10.1115/GT2016-56561.

In this paper, we present an extensive numerical study on the interaction between the downstream fan and the flow separating over an intake under high incidence. The objectives of this investigation are twofold: (a) to gain qualitative insight into the mechanism of fan-intake interaction and (b) to quantitatively examine the sensitivity of the flow distortion (in terms of distortion coefficient DC60), to the key design parameters of the intake (Length, L / Diameter, D).

Both steady and unsteady Reynolds Averaged Numerical Simulations (RANS) were carried out. For the steady calculations, a low order fan model has been used while a full 3D geometry has been used for the unsteady RANS. The numerical methodology is also thoroughly validated against the measurements for the intake-only and fan-only configurations on a high bypass ratio turbofan intake and fan respectively. To systematically study the effect of fan on the intake separation and explore the design criteria, a simplified intake-fan configuration has been considered. In this fan-intake model, the ratio of the intake length to diameter (L/D) can be conveniently altered without affecting other parameters.

The key results indicate that, depending on L/D, the fan has either suppressed the level of the post separation distortion or increased the separation-free operating range. At the lowest L/D (∼ 0.17), around a 5° increase in the separation-free angle of incidence was achieved. This delay in the separation-free angle of incidence decreased with increasing L/D. At the largest L/D (∼ 0.44), the fan was effective in suppressing the post-separation distortion rather than entirely eliminating the separation. Isentropic Mach number distributions over the intake lip for different L/D’s revealed that the fan accelerates the flow upstream of the fan face, thereby decreasing the distortion level in the immediate vicinity. However, this acceleration effect decayed rapidly with increasing upstream distance from the fan-face.

Commentary by Dr. Valentin Fuster
2016;():V001T01A015. doi:10.1115/GT2016-56572.

This paper describes research carried out in the European Commission co-funded project E-BREAK (Engine BREAK through components and subsystems) focused on development of generic enabling technologies for new aero-engines.

A global market forecast (2015–2034) from Airbus [1], depicts an average growth rate of 4.6% per year. Air traffic is forecasted to double in the next 15 years. It is expected, to triple in the next 20 years, according to the speech given by RRUK CEO during the Aerodays 2015 in London [2]. This high level of growth in demand for air travel represents huge opportunities as well as significant challenges for the aerospace industry. Research and Technology through collaborative European projects addresses the environmental penalties of air traffic. Europe’s aviation industry therefore faces a huge challenge to satisfy the demand whilst guaranteeing competitiveness, safety and more environmentally friendly air travel. Innovative engine configurations consequently need to be investigated in order to reduce significantly the pollutant emissions (15 to 20% for fuel consumption and CO2 and 80% reduction for NOx). Such reductions can only be achieved by considering innovative components that could be integrated and optimized in new engine configurations.

In response to the above demands, aero-engine manufacturers are constantly aiming to improve gas turbine efficiency for two main reasons: to reduce environmental impact and to minimize operating costs.

The E-BREAK project is aimed at the development of generic enabling technologies needed to address the challenges for future engines with higher overall pressure ratios (OPR) and bypass ratio (BPR). These technologies are developed at subsystem and component level and validated in test rigs which are equivalent to Technical Readiness Level (TRL) 5. The utility of the developed technologies are assessed using four standard study powerplants. These are turboshaft, regional turbofan, mid-size open rotor, and large turbofan, covering most of the expected future commercial aero-engine market.

This article describes the technical approach followed in E-BREAK for the various technologies being investigated, these are:

• Advanced sealing to reduce oil and air leakages

• Variability control to ensure stability of thermodynamic cycle

• High temperature resistant material and abradables to prevent fast degradation at high temperatures

• Light material to prevent significant mass increase

• Health monitoring system to anticipate sub-systems degradation

The envisaged outcomes from E-BREAK are enablers to other EU-funded research projects focused on module maturation progress, such as LEMCOTEC dealing with high OPR modules and ENOVAL dealing with high BPR LP components.

Commentary by Dr. Valentin Fuster
2016;():V001T01A016. doi:10.1115/GT2016-56582.

The jetted hot film could affect the trajectories of the water droplets near the aero-engine inlet strut surface, which equipped with the ice protection system combined the internal impingement heat transfer and the external hot air film heating. To evaluate the droplets impingement characteristics of four ice-protection structures designed with different film-slot jet angles, a droplets impingement computation method based on Eulerian framework was developed and validated. The influences of film-slot angle and blowing ratio on the impingement characteristics for droplet diameter of 20μm were investigated and the jet vortex was found to be an important factor. The results indicated that the local collection efficiency and the impingement limitation could decrease significantly due to the blowing from the external hot-film, and the influence would be more significant in case that the film-slot was closer to the leading edge. For example, the average local collection efficiencies of four typical configurations with different slot angles and positions decreased 82%, 8%, 1% and 0.5% respectively comparing to those without air film. Besides that, the maximum local collection efficiency and the impingement limitation decreased with the increasing blowing ratios, and the film-slot nearest to the leading edge was most sensitive to the blowing ratio. It was also found that no droplets impinged on the rear surface after the jet slots at some higher blowing ratios.

Commentary by Dr. Valentin Fuster
2016;():V001T01A017. doi:10.1115/GT2016-56633.

The development and limitations of a numerical modelling framework applied to an aero-engine air/oil separator are presented here. Oil enters the device in the form of dispersed droplets and primary separation occurs by centrifuging larger droplets towards the outer walls, whereas secondary separation occurs by partially coalescing and centrifuging smaller droplets within a porous material, namely an open-cell metal foam. The work described here is part of a study led jointly by the University of Nottingham (UNott) and the Karlsruhe Institute of Technology (KIT) in the Engine Breakthrough Components and Subsystems (E-BREAK) project. The main objectives for UNott have been to define a CFD methodology able to provide an accurate representation of the air flow behaviour and a qualitative assessment of the oil capture within the air/oil separator. The feasibility of using the current state-of-the-art modelling framework is assessed. Experimental measurements of the overall pressure drop and oil capture performed at KIT are used to validate the simulations. The methodology presented here overcomes some limitations and simplifications present in previous studies. A novel macroscopic model for the secondary oil separation phenomena within metal foams is presented. Experiments and simulations were conducted for three different separator configurations, one without a metal foam, and two with metal foams of different pore sizes. For each configuration, a variation of air flow, shaft speed and droplet size was conducted. The focus was on the separation of droplets with a diameter smaller than 10 μm. Single-phase air flow simulation results showed that overall pressure drop increases with both increased shaft speed and air flow, largely in agreement with the experiments. Oil capture results proved to be more difficult to be captured by the numerical model. One of the limitations of the modelling set-up employed here is not capable of capturing droplet re-entrainment due to accumulation of oil inside the metal foam, which is believed to play a significant role in the separation phenomena.

Commentary by Dr. Valentin Fuster
2016;():V001T01A018. doi:10.1115/GT2016-56645.

This paper presents a method for modelling contra-rotating propellers (CRP) for engine performance simulations. An in-house free-wake lifting surface tool (GENUVP) is used to generate suitable performance maps for each propeller that express power and thrust coefficient in terms of advance ratio, flight Mach number, speed ratio and blade pitch angle of each propeller. Appropriate component models that utilize these maps are then developed in a commercial engine performance simulation environment (PROOSIS). Next, the propeller components are integrated in a direct-drive open rotor engine model. Finally, design point and off-design simulations are carried out that demonstrate the use of the model through studies of different propeller blade angle control strategies.

Commentary by Dr. Valentin Fuster
2016;():V001T01A019. doi:10.1115/GT2016-56708.

This paper presents the design and full-scale ground-test demonstration of an engine air-brake (EAB) nozzle that uses a deployable swirl vane mechanism to switch the operation of a turbofan’s exhaust stream from thrust generation to drag generation during the approach and/or descent phase of flight. The EAB generates a swirling outflow from the turbofan exhaust nozzle, allowing an aircraft to generate equivalent drag in the form of thrust reduction at a fixed fan rotor speed. The drag generated by the swirling exhaust flow is sustained by the strong radial pressure gradient created by the EAB swirl vanes. Such drag-on-demand is an enabler to operational benefits such as slower, steeper, and/or aeroacoustically cleaner flight on approach, addressing the aviation community’s need for active and passive control of aeroacoustic noise sources and access to confined airports.

Using NASA’s Technology Readiness Level (TRL) definitions, the EAB technology has been matured to a level of 6, i.e., a fully functional prototype. The TRL-maturation effort involved design, fabrication, assembly, and ground-testing of the EAB’s deployable mechanism on a full-scale, mixed-exhaust, medium-bypass-ratio business jet engine (Williams International FJ44-4A) operating at the upper end of typical approach throttle settings. The final prototype design satisfied a set of critical technology demonstration requirements that included (1) aerodynamic equivalent drag production equal to 15% of nominal gross thrust in a high-powered approach throttle setting (called dirty approach), (2) excess nozzle flow capacity and fuel burn reduction in the fully deployed configuration, (3) acceptable engine operability during dynamic deployment and stowing, (4) deployment time of 3–5 seconds, (5) stowing time under 0.5 second, and (6) packaging of the mechanism within a notional engine cowl. For a typical twin-jet aircraft application, a constant-speed, steep approach analysis suggests that the EAB drag could be used without additional external airframe drag to increase the conventional glideslope from 3 to 4.3 degrees, with about 3 dB noise reduction at a fixed observer location.

Commentary by Dr. Valentin Fuster
2016;():V001T01A020. doi:10.1115/GT2016-56747.

This paper presents a coupled ETFM-VOF framework for the numerical simulation of multi-scale thin liquid films. A depth-averaged Eulerian thin-film model (ETFM) is used to simulate the oil flow in very thin-film regions where film thicknesses are below the grid resolution while elsewhere in the domain where grid resolution is sufficient to resolve the film, a traditional Volume-of-Fluid (VOF) approach is retained. The two approaches are coupled through momentum and mass conserving source terms and a transition criterion is introduced where the total liquid volume fraction in each cell is evaluated and either the ETFM or VOF approach used depending on the sufficiency of the local grid resolution. Using this approach, thin-film flows characterised by multiple film thickness scales may be reliably simulated at a relatively lower computational cost. The model builds upon currently available ETFM and VOF approaches to thin-film modelling and represents a novel approach to the numerical simulation of multiphase flows involving a varying range of film thickness scales in space and time. A numerical test case of the 3D rimming flow inside an idealised aero-engine bearing chamber has been used to demonstrate the approach and comparisons made against high resolution VOF solutions.

Commentary by Dr. Valentin Fuster
2016;():V001T01A021. doi:10.1115/GT2016-56923.

In this paper, a cost-effective integrated simulation model for rotor/propeller driven aerobot is proposed based on the free wake model. It is helpful to provide aerodynamic input for preliminary control law design, especially for conceptual design at the beginning of a project. Apart from the computational efficient, the model proposed has two more advantages, grid-free and containing the aerodynamic interaction. The governing equation is the Laplacian Equation with the assumption of invisid, incompressible flow. The solver used is inspired by the well-known Pseudo-Implicit Predictor Corrector (PIPC) for the rotors. Two validation cases are carried out. Firstly, the simulation results for the Harrington coaxial rotors and the Hamilton coaxial propellers show good agreement with the experiment. The simulation of a 45° swept-back wing also goes well with the corresponding experiment. Before the end, a simulation for a transition state with 0° tilt angle for the Eagle Eye is used to show the aerodynamic interaction, which has more influence on the wings.

Commentary by Dr. Valentin Fuster
2016;():V001T01A022. doi:10.1115/GT2016-57074.

Numerical simulation of a turbojet engine with water injection from the compressor inlet has been executed to evaluate the effect of wet compression on the pollutant emission in this paper. Water injection technique has shown the advantage of reducing the compression work and increasing the power output, and can also reduce the gas temperature in the combustion chamber and turbine which can help to reduce the thermal fatigue of the hot parts and extend their service life. Previous studies have indicated that there is a great relationship between pollutant emission of NOx and the combustion temperature, so the decrease of combustion temperature can reduce the amount of pollutant production of aircrafts especially near the airport because of the takeoff and landing of the aircrafts. And this potential of meeting the increasingly stringent pollutant emission target of water injection technique has aroused great interest of the world major engine manufacturers and research institutions. Compared with cases of no water injection, the results of this study show that the inlet temperature of combustion chamber drops obviously because of water injection, the average NO concentration in combustion chamber decreases greatly, but the concentration of CO increases significantly. Analysis of the results shows that based on fuel consumption of per unit mass the relation between the NO production in combustion chamber and the mass water evaporation varies according to the exponential law and CO production in water injection cases is similar to the cases without water injection.

Commentary by Dr. Valentin Fuster
2016;():V001T01A023. doi:10.1115/GT2016-57086.

A new nose cone ice protection configuration with hot air film slot was investigated, where the surface need to be protected could be heated with interior impingement heat transfer and exterior hot air film. Numerical simulation methods using computational fluid dynamics code were developed and validated to find the effects of the jetted air film on the droplet impingement characteristics. Combination of two-dimensional axisymmetric algorithm and Lagrangian method were adopted to solve the air flow field and the droplet trajectories. The simulation methods were validated with the results from the experimental data. The droplet impingement characteristics on two structures were investigated respectively, the intact cone without film slot and the slotted cone. Results show that the surface local collection coefficient changes significantly behind the film slot because of the blowing effect of the air film on the incoming droplets, and the variation of the local collection coefficient is very dependent on the droplet diameter for a given blowing ratio, or the blowing ratio for a given droplet size. Some different effects, such as “fully blowing-off”, “blowing behind” and “limited blowing-off”, may happen for different combination of droplets size and blowing ratio. Compared with the cone without film slot, the blowing effect is more significant for smaller droplets or higher blowing ratio. Besides that, the total collection coefficient maybe only half of those without film for some conditions.

Commentary by Dr. Valentin Fuster
2016;():V001T01A024. doi:10.1115/GT2016-57257.

Transient effects are important features of engine performance calculations. The aim of this paper is to analyze a new, fully transient model implemented using the PRopulsion Object Oriented Simulation Software (PROOSIS) for a civil, short range turbofan engine. A transient turbofan model, including the mechanical inertia effect has been developed in PROOSIS. Specific physical effects such as heat soakage, mass storage, blade tip clearance and combustion delay have been implemented in the relevant components of PROOSIS to obtain a fully transient model. Since a large number of components are concerned by all the transient effects, an influence study is presented to determine which are the most critical effects, and in which components. Inertia represents the relevant phenomenon, followed by thermal effects, combustion delay and finally mass storage. The comparison with experimental data will provide a first validation of the model. Finally a sensitivity study is reported to assess the impact of uncertain knowledge of key input parameters in the response time prediction accuracy.

Commentary by Dr. Valentin Fuster
2016;():V001T01A025. doi:10.1115/GT2016-57307.

This paper describes the research carried out in the European Commission co-funded project LEMCOTEC (Low Emission Core Engine Technology) on aerodynamics for turbines and structures for compressors, combustors and turbines. The aim is to significantly contribute to the reduction of the environmental footprint of aviation with regard to emissions from aero engines.

The LEMCOTEC turbine and structure technologies are directed primarily to act as enablers for higher thermal efficiency arising from increased overall pressure ratio. Thus the work is supporting increased operating temperatures, reduced core deformation, reduced cooling flows and increased performance to weight ratio, in addition to direct reduction of flow losses and associated component efficiency increases.

The article details the targets for performance improvements, the validation of the technologies and how they, together with LEMCOTEC’s improved technologies on compressors and combustors, relate to the goal of building ultra-high pressure ratio engines.

Commentary by Dr. Valentin Fuster
2016;():V001T01A026. doi:10.1115/GT2016-57332.

Threats to engine integrity and life from deposition of environmental particulates that can reach the turbine cooling systems (i.e. <10 micron) have become increasing important within the aero-engine industry, with an increase of flight paths crossing sandy, tropical storm-infested, or polluted airspaces. This has led to studies in the turbomachinery community investigating environmental particulate deposition, largely applying the Discrete Random Walk (DRW) model in CFD simulations of air paths. However, this model was conceived to model droplet dispersion in bulk flow regimes, and therefore has fundamental limitations for deposition studies. One significant limitation is an insensitivity to particle size in the turbulent deposition size regime, where deposition is strongly linked to particle size. This is highlighted within this study through comparisons to published experimental data.

Progress made within the wider particulate deposition community has recently led to the development and application of the Continuous Random Walk (CRW) model. This new model provides significantly improved predictions of particle deposition seen experimentally in comparison to the DRW for low temperature pipe flow experiments. However, the CRW model is not without its difficulties. This paper highlights the sensitivities within the CRW model and actions taken to alleviate them where possible. For validation of the model at gas turbine conditions, it should be assessed at engine-representative conditions. These include high-temperature and swirling flows, with thermophoretic and wall-roughness effects. Thermophoresis is a particle force experienced in the negative direction of the temperature gradient, and can strongly effect deposition efficiency from certain flows. Previous validation of the model has centred on low temperatures and pipe flow conditions. Presented here is the validation process which is currently being undertaken to assess the model at gas turbine-relevant conditions. Discussion centres on the underlying principles of the model, how to apply this model appropriately to gas turbine flows and initial assessment for flows seen in secondary air systems. Verification of model assumptions is undertaken, including demonstrating that the effect of boundary layer modelling of anisotropic turbulence is shown to be Reynolds-independent. The integration time step for numerical solution of the non-dimensional Langevin equation is redefined, showing improvement against existing definitions for the available low temperature pipe flow data. The grid dependence of particle deposition in numerical simulations is presented and shown to be more significant for particle conditions in the diffusional deposition regime. Finally, the model is applied to an engine-representative geometry to demonstrate the improvement in sensitivity to particle size that the CRW offers over the DRW for wall-bounded flows.

Commentary by Dr. Valentin Fuster
2016;():V001T01A027. doi:10.1115/GT2016-57463.

A correctly profiled engine nacelle can delay the transition in the boundary layer and allow laminar flow to extend back, resulting in a substantial drag reduction. Therefore, the laminar flow nacelle has lower fuel consumption than current turbulent designs. In this paper, aerodynamic shape optimization of natural laminar flow nacelle has been studied by using a novel nacelle shape design method and transition prediction with CFD. First, the 2D longitudinal profile-line of nacelle is optimized, in order to extend its laminar region and achieve minimum drag coefficient within the design space. Second, the optimized longitudinal profile-line is then circumferentially stacked to construct the 3D nacelle aerodynamic shape. At last, the aerodynamic improvement of the new shape is evaluated by 3D CFD simulation. A nacelle geometry generator has been developed where the deflection angle (related to the curvature) along the cord is controlled by using Non-Uniform Rational B-Splines. It is then analytically integrated to obtain the longitudinal profile-line. And also a leading edge matching function is involved in the generator. This technique improves the smoothness of nacelle profile-line, which ensures the curvature and slope of curvature to be continuous all over the nacelle surface. The pressure distribution over the nacelle surface has been improved with no spikes in Mach number. A transition model coupling with shear stress transport turbulent model is used in solving Navier-Stokes equations for transition prediction. An optimization system has been established in combination with the geometry generator, the transition prediction model with CFD, a Kriging surrogate model and a Multi-Island Genetic Algorithm. As a result, the aerodynamic improvement, with one profile-line optimized, is obvious against the original nacelle shape by CFD validation in 3D simulation. The optimized nacelle can achieve a laminar flow up to 23% and its drag coefficient has reduced by 6.5%. It is indicated that the optimization system is applicable in nacelle aerodynamic shape design.

Commentary by Dr. Valentin Fuster
2016;():V001T01A028. doi:10.1115/GT2016-57524.

This paper investigates analytically the advantage of the embedded propulsion compared to a state of the art propulsion of an aircraft. Hereby, we are applying the integral method of boundary layer theory and potential theory to analyse the boundary layer thickness and the impact of the flow acceleration due to the embedded propulsion. The aircraft body is treated as a flat plate. The engine is treated as a momentum disc but there is a trade off, since the engine efficiency is effected by the boundary layer. The outcome of the energetic assessment is the following: the propulsion efficiency is increased by the embedded propulsion and the drag of the aircraft body is reduced. The optimized aircraft engine size depending on Reynolds number is given.

Commentary by Dr. Valentin Fuster
2016;():V001T01A029. doi:10.1115/GT2016-57539.

Motivated by the long term target settings for research and innovation in Europe and in North America, initial investigations of parallel hybrid electric power plant systems have indicated significant fuel reduction potentials for short range air transport. While an electric motor assists the gas turbine in suppling mechanical power to the gas turbine within the classical parallel hybrid topology, in the present paper, a more sophisticated variant, namely the Cycle-Integrated Parallel Hybrid (CIPH) is considered. More specifically, the design and performance implications of a CIPH power plant architecture are investigated with regard to an advanced turboshaft (TS) engine application for helicopters. For this purpose, an array of compressor stages of the baseline TS power plant are decoupled from the turbine section, and are driven mechanically independently by means of electric motors. The baseline power plant of the investigated concept is derived for a 12-ton-helicopter accommodating 19 passengers on a 450nm mission. It consists of an axi-centrifugal compressor powered by the high pressure turbine as well as a free low pressure (power) turbine delivering a maximum shaft power of 3300 kW. For the presented CIPH concept, the axial compressor section is electrified with the help of linear electric motors mounted at the blade tips. Due to typical design characteristics of electric motors, counter rotating stages are considered most appropriate for the targeted TS power plant application. The electric motor power supply is realized through a Power Management And Distribution system featuring proper levels of redundancy. For the electrical energy storage, advanced battery technology is taken into account. Hybrid electric Energy and Propulsion Systems (EPS) can be characterized meaningfully by the degree of power hybridization, HP, being defined as the ratio of the installed electric power to the total power. For the presented CIPH application, a best and balanced HP of 19.7% has been identified. In typical part load operation, this may lead to relative Power Specific Fuel Consumption (PSFC) improvements of up to 45% and overall efficiency has been almost doubled compared to the TS reference. With the implementation of electric power within the cycle, additional degrees of freedom for power plant operation and control can be established. At vehicular level, a retrofitted version of the reference helicopter equipped with the CIPH TS propulsion system faces more than 50% reduced range, but simultaneously reduces total energy consumption (fuel plus electrical) by 28% and CO2 by 42% compared to the reference vehicle at identical reduced range.

Topics: Cycles
Commentary by Dr. Valentin Fuster
2016;():V001T01A030. doi:10.1115/GT2016-57691.

Compared with the single propeller, a propeller with Swirl Recovery Vanes (SRV) has great advantages in recovering the residual swirl, which leads to an improvement both in thrust and propulsion efficiency. Based on the preliminary design of the SRV for Fokker F29 propeller, Design of Experiment (DoE) method was employed to optimize and further improve the performance of SRV in the present study. Firstly, orthogonal experiment was adopted to identify the most significant factors for affecting thrust. Secondly, steepest ascent method was used to search for the optimum range of target factors by using climbing experiment and factorial experiment. Finally, center composite experiment was adopted to find the optimal solution. As a result, the thrust of the optimized SRV has increased significantly by 11.78% at the design point, which leads to a 0.66% increment of total efficiency of propeller and SRV propulsion system, and even more increment at low rotation speed (2.10%). Besides, compared with the original SRV, the optimized SRV geometry improves the flow characteristics at the tip region of the vane. The tip vortex and swirl kinetic energy have been weakened distinctly, and the thrust coefficient distribution along the spanwise tends to be more uniform.

Commentary by Dr. Valentin Fuster
2016;():V001T01A031. doi:10.1115/GT2016-57719.

There is a high potential for civil applications of Unmanned Aerial Vehicles (UAV) in areas such as goods transport, telecommunication, remote monitoring and sensing, surveillance, search and rescue, and disaster management. Developments in areas such as telecommunication, control and information technology offer opportunities for long range remotely or automatically piloted missions. This requires efficient and light-weight small propulsion systems.

The potential of turboprop propulsion for civil UAVs using micro turbine technology has been explored and compared with existing concepts, such as piston engine driven propellers. Different propulsion concepts have been analyzed and the application areas where advanced turboprops would be superior to other systems such as reciprocating engines and electric motors, identified. However, turboprop engines of the small power capacity required for the aircraft concepts and missions considered are not currently available with competitive performance.

A conceptual design study of a micro turboprop engine has been performed by downscaling an existing reference engine. Scale effects on efficiency have been taken into account, as well as effects of technological progress. Engine cycle optimization has been carried out and the effects of turbine inlet temperature, compressor pressure ratio, engine size, and component efficiency have been investigated. An aerodynamic and flight performance model of a baseline UAV has been developed in order to predict mission performance. This model has been coupled to a turboprop model to evaluate system performance with different engine configurations for the selected mission.

The outcome of the study provides information about the technological improvements in terms of cycle efficiency required to make the micro-turboprop a competitive solution. The Propulsion and Power group of Delft University of Technology will pursue these R&D goals in an attempt to contribute to the development of civil UAV technology.

Commentary by Dr. Valentin Fuster
2016;():V001T01A032. doi:10.1115/GT2016-57724.

Cruise specific fuel consumption (SFC) of turbofan engines is a key metric for increasing airline profitability and for reducing CO2 emissions. Although increasing design bypass ratio (BPR) of separate exhaust turbofan configurations improves cruise SFC, further improvements can be obtained with control actuated variable geometry modulations of core nozzle throat area, bypass nozzle throat area, and compressor variable vanes (CVV). The scope of this paper is to show only the benefits possible, and the process used in determining those benefits, and not to suggest any particular control algorithm for searching the best combination of the control effectors. A parametric cycle study indicated that the effector modulations could increase the cruise BPR, core efficiency, transmission efficiency, propulsive efficiency, and ideal velocity ratio resulting in a cruise SFC improvement of as much as 2.6% depending upon the engine configuration. The changes in these metrics with control effector variations will be presented. Modulation of CVV is already possible in legacy digital controls, and modulation of nozzle areas should be explored in light of the low bandwidth requirements at steady-state cruise conditions.

Commentary by Dr. Valentin Fuster
2016;():V001T01A033. doi:10.1115/GT2016-57784.

A parametric geometry definition for a generic turbofan nacelle was developed for use in preliminary design, based on Class-Shape Transformation curves. This takes as input a set of six intuitive variables which describe the main dimensions of a nacelle. This set is the same set of inputs as required by a preliminary nacelle design method to which the aerodynamic properties of resulting shapes were compared. An automated computational fluid simulation process was developed and implemented which generates meshes and quickly conducts an analysis of the resulting nacelle shapes using a commercial code. Several geometries were generated and analysed using this process to show whether the aerodynamic properties of the generated shapes are in line with the expected performance of a fan cowl of equal dimensions. It was found that the aerodynamic performance of the parametric fan cowls significantly exceeds predictions from an established preliminary fan cowl design method and is very close in performance to existing designs. The drag of an equivalent parametric fan cowl can therefore be used as a predictor of nacelle performance with greater accuracy than established preliminary design methods. It is therefore suited as a tool to develop improved preliminary design methods, and for studies of the design space for preliminary nacelle design.

Commentary by Dr. Valentin Fuster
2016;():V001T01A034. doi:10.1115/GT2016-57849.

A tool to create parametric aerodynamic shapes using intuitive design variables based on class shape transformation curves is presented. To enable this, a system has been developed which accepts arbitrary constraints and automatically derives the analytical expressions which describe the corresponding class shape transformation curves. Parametric geometry definitions for fan cowl and intake aero-lines were developed using the generalized method. CFD analysis of the fan cowl shows that despite the simple geometry definition its performance characteristics are close to what would be expected of a finished design. The intake geometry was generated in a similar way and met the typical performance metrics for conventional intakes. This demonstrates the usefulness of the tool to quickly and robustly produce parametric aero-lines with good aerodynamic properties, using relatively simple intuitive design variables.

Topics: Shapes
Commentary by Dr. Valentin Fuster
2016;():V001T01A035. doi:10.1115/GT2016-57854.

Although development of the line of gas turbine engines on the basis of a unified engine core is a widely used practice, the method of selection of the most efficient values of engine core parameters has virtually never been published.

The paper describes the method of optimization of values of engine core parameters jointly with optimization of values of every engine forming the line of engines developed on the basis of this core. Both stages of optimization are multi-objective and fulfilled using the results of simulation of engine as a subsystem of the airframe.

Commentary by Dr. Valentin Fuster
2016;():V001T01A036. doi:10.1115/GT2016-57875.

This paper reports the internal performance evaluation of S-duct diffusers with different entrance aspect ratios as part of an ongoing parametric investigation of a generic S-duct inlet. The generic S-duct diffusers were a rectangular-entrance (aspect ratio 1.5 and 2.0) transitioning S-duct diffuser in high subsonic (Mach number > 0.8) flow. The test section was manufactured using rapid prototyping for facilitating the parametric investigation of the geometry. Streamwise static pressure and exit-plane total pressure were measured in a test-rig using surface pressure taps and a 5-probe rotating rake, respectively and the baseline and a variant was simulated through computational fluid dynamics. The investigation indicated the presence of streamwise and circumferential pressure gradients leading to a three dimensional flow in the S-duct diffuser and distortion at the exit plane. The static pressure recovery increased for the diffuser with higher aspect ratio. Total pressure losses and circumferential and radial distortions at the exit plane were higher than that of the podded nacelle type of inlet. The increase in the total pressure recovery was observed for the increase in the aspect ratio for the baseline area ratio (1.57) S-ducts, but without a clear trend for the other area ratio (1.8) ducts. The work represents the beginning of the development of a database for the performance of a particular type of generic inlet. This database will be useful for predicting the performance of aero-engines and air vehicles in high subsonic flight.

Commentary by Dr. Valentin Fuster
2016;():V001T01A037. doi:10.1115/GT2016-57918.

Direct Drive Open Rotors (DDORs) have the potential to significantly reduce fuel consumption and emissions relative to conventional turbofans. However, this engine architecture presents many design and operational challenges both at engine and aircraft level. At preliminary design stages, a broad design space exploration is required to identify potential optimum design regions and to understand the main trade offs of this novel engine architecture. These assessments may also aid the development process when compromises need to be performed as a consequence of design, operational or regulatory constraints.

Design space exploration assessments are done with 0-D or 1-D models for computational purposes. These simplified 0-D and 1-D models have to capture the impact of the independent variation of the main design and control variables of the engine. Historically, it appears that for preliminary design studies of DDORS, Counter Rotating Turbines (CRTs) have been modeled as conventional turbines and therefore it was not possible to assess the impact of the variation of the number of stages (Nb) and rotational speed of the propellers. Additionally, no preliminary design methodology for CRTs was found in the public domain.

Part I of this two-part publication proposes a 1-D preliminary design methodology for DDOR CRTs. It allows an independent definition of the Nb, rotational speeds of both parts of the CRT, inlet flow conditions, inlet and outlet annulus geometry as well as power extraction. It includes criteria and procedures to calculate: power extraction in each stage, gas path geometry, blade metal angles, flow conditions at each turbine plane and overall CRT efficiency. The feasible torque ratios of a CRT are discussed in this paper. A form factor for the CRT velocity triangles is defined (similar to stage reaction on conventional turbines) and its impact on performance and blade design is discussed. A method for calculating the off-design performance of a CRT is also described in Part I.

In Part II, a 0-D design point (DP) efficiency calculation for CRTs is proposed as well as a case study of a DDOR for a 160 PAX aircraft. In the case study, three main aspects are investigated: A) the design and performance of a 20 stage CRT for the DDOR application; B) the impact of the control of the propellers on cruise specific fuel consumption, C) the impact of the design rotational speeds and Nb of the CRT on its DP efficiency, engine fuel consumption and engine weight.

Topics: Design , Rotors , Turbines
Commentary by Dr. Valentin Fuster
2016;():V001T01A038. doi:10.1115/GT2016-57921.

Direct Drive Open Rotors (DDORs) have the potential to significantly reduce fuel consumption and emissions relative to conventional turbofans. However, this engine architecture presents many design and operational challenges both at engine and aircraft level. At preliminary design stages, a broad design space exploration is required to identify potential optimum design regions and to understand the main trade offs of this novel engine architecture. These assessments may also aid the development process when compromises need to be performed as a consequence of design, operational or regulatory constraints.

Design space exploration assessments are done with 0-D or 1-D models for computational purposes. These simplified 0-D and 1-D models have to capture the impact of the independent variation of the main design and control variables of the engine. Historically, it appears that for preliminary design studies of DDORs, Counter Rotating Turbines (CRTs) have been modelled as conventional turbines and therefore it was not possible to assess the impact of the variation of the number of stages (Nb) of the CRT and rotational speed of the propellers. Additionally, no preliminary design methodology for CRTs was found in the public domain.

Part I of this two-part publication proposes a 1-D preliminary design methodology for DDOR CRTs which allows an independent definition of both parts of the CRT. A method for calculating the off-design performance of a known CRT design is also described.

In Part II, a 0-D design point efficiency calculation for CRTs is proposed and verified with the 1-D methods. The 1-D and 0-D CRT models were used in an engine control and design space exploration case study of a DDOR with a 4.26m diameter an 10% clipped propeller for a 160 PAX aircraft. For this application:

• the design and performance of a 20 stage CRT rotating at 860 rpm (both drums) obtained with the 1-D methods is presented.

• differently from geared open rotors, negligible cruise fuel savings can be achieved by an advanced propeller control.

• for rotational speeds between 750 and 880 rpm (relatively low speeds for reduced noise), 22 and 20 stages CRTs are required.

• engine weight can be kept constant for different design rotational speeds by using the minimum required Nb.

• for any target engine weight, TOC and cruise SFC are reduced by reducing the rotational speeds and increasing Nb (also favourable for reducing CRP noise). However additional CRT stages increase engine drag, mechanical complexity and cost.

Topics: Design , Rotors , Turbines , Aircraft
Commentary by Dr. Valentin Fuster
2016;():V001T01A039. doi:10.1115/GT2016-58012.

This paper proposes a preliminary subsonic aircraft and engine noise assessment framework, capable of computing the aircraft total noise level at all three certification points (i.e. Approach, Lateral, and Flyover) defined by the International Civil Aviation Organisation. The proposed framework is numerically integrated to account for the complete aircraft noise sources (i.e. the fuselage, wings, landing gear, as well as noise sources resulting from the engine component level, (i.e. fan, compressor, combustor, turbine, and jet). The developed framework is based on a wide-range of empirical and semi-empirical correlations collected from the public domain literature. The fidelity of the framework also caters for flight effects such as atmospheric attenuation, spherical spreading, Doppler shift, lateral attenuation, retarded time and ground reflection. A conversion between the sound pressure level SPL [SPLdB] to effective perceived noise level EPNL [EPNdB] is also included to allow for a consistent comparison with the certification procedure. Through the successful deployment of the proposed framework a generic aircraft model, representative of a modern commercial carrier aircraft has been investigated, operating under representative operational conditions. The sound pressure level corresponding to various aircraft and engine component have been thoroughly investigated and verified with trends acquired based on the theory. Furthermore, the predictions made by the framework corresponding to the aforementioned three certification points have also been verified against the noise level measurements provided by the International Civil Aviation Organization. The results acquired exhibit good correlation against the verification data for total noise levels at the microphones. Furthermore, a component level comparison is also presented which exhibit good agreement with verification data. The deployed methodology can essentially be regarded as an enabling technology to support the effective and efficient implementation of framework(s) (i.e. Technoeconomic, Environmental and Risk Assessment) targeted to evaluate the existing and advanced aircraft and engine architectures in terms of operational performance and environmental impact.

Commentary by Dr. Valentin Fuster

Fans and Blowers

2016;():V001T09A001. doi:10.1115/GT2016-56211.

Jet fans produce a highly anisotropic turbulent flow regime and cause severe entrainment of the ambient fluid. Additionally, the placement of jet fans near or away from walls has a significant impact on throw and spread of the fan. This paper proposes a RSM based numerical technique to accurately predict the throw and spread of jet fans in an industrial setting while accounting for both the near wall effects and anisotropic nature of turbulence. The RSM based numerical technique described in this study allows for accurate prediction of the axial throw termination distances for jet fans in an industrial setting. The RSM approach also overcomes the limitations imposed by the free jet theory by accounting for wall effect on flow as well as other physical or temporal limitations imposed by various numerical methods such as RANS and SRS.

The Free Jet Theory proposes a classification method for turbulent jet decay problems, where the jet decay domain is classified into three flow domains and one terminal region. The terminal region of turbulent jet decay is not understood well physically and is neglected for the purpose of throw and spread calculations. The technique put forth in this paper draws from the fundamentals presented by the free jet theory and is then modified and applied to account for Coanda effects seen due to the proximity of walls to the flow when the jet fan is mounted close to the ceiling. Finally, experimental data is recorded by placing the jet fan in proximity with the floor in a closed room. The data thus generated is subsequently used to validate numerical results.

Commentary by Dr. Valentin Fuster
2016;():V001T09A002. doi:10.1115/GT2016-56299.

The present paper focuses on the potential efficiency improvements and the stable operating range of a centrifugal fan for fuel cell applications. Improvements will be achieved by variability of the cross-sectional area of diffuser and volute by use of a moving backplate.

The investigation consists of three parts: The first part describes the design and the performance prediction of a diffuser-volute combination with a variable cross-sectional area, based on empirical correlations and low-resolution CFD (Computational Fluid Dynamics) simulations. For the second part, high-resolution 360 degree CFD simulations are used to gain deeper insight into the flow mechanisms and their influence on fan performance. The last part presents the experimental investigations carried out to validate the numerical models. For this purpose, a demonstrator of the fan including a diffuser-volute combination with variable cross-sectional area is manufactured and investigated using optical PIV (Particle Image Velocimetry) measurements.

Topics: Diffusers , Fans
Commentary by Dr. Valentin Fuster
2016;():V001T09A003. doi:10.1115/GT2016-56491.

The performance of large mechanical draft air-cooled heat exchangers is directly related to fan performance which is influenced by atmospheric wind conditions, as well as the plant layout. It is often necessary to numerically model the entire system, including fans, under a variety of operating conditions.

Full three-dimensional, numerical models of axial flow fans are computationally expensive to solve. Simplified models that accurately predict fan performance at a lesser expense are therefore required. One such simplified model is the actuator disk model (ADM). This model approximates the fan as a disk where the forces generated by the blades are calculated and translated into momentum sources. This model has been proven to give good results near and above the design flow rate of a fan, but not at low flow rates. In order to address this problem two modifications were proposed, namely the extended actuator disk model (EADM) and the reverse engineered empirical actuator disk model (REEADM).

The three models are presented and evaluated in this paper using ANSYS FLUENT. The models are simulated at different flow rates representing an axial flow fan test facility. The resulting performance results and velocity fields are compared to each other and to previously simulated three dimensional numerical results, indicating the accuracy of each method. The results show that the REEADM gives the best correlation with experimental performance results at design conditions (ϕ = 0.168) while the EADM gives the best correlation at low flow rates.

A comparison of the velocity profiles shows that none of the three models predict the radial velocity distribution at low flow rates correctly, however the correlation improves at flow rates above ϕ = 0.105. In general the upstream velocity profiles, where reversed flow occurs through the fan, are poorly predicted at low flow rates. At the flow rates above ϕ = 0.137 the correlation between the velocity profiles for the simplified modes and the three dimensional results is good.

Commentary by Dr. Valentin Fuster
2016;():V001T09A004. doi:10.1115/GT2016-56555.

Not only the aerodynamic performance of axial flow fans is important but also the acoustic behaviour plays a vital role. It is to be expected that in the future noise limits will be more regulated by legislation. The aim of this project is to develop a very versatile tool for efficient and noise reduced axial flow fans in rotor / stator configuration.

This paper describes the design, numerical verification and tests of a highly loaded single stage axial flow fan making use of extensive blade sweep in rotor and stator for acoustic reasons. The tests include aerodynamic and acoustic investigations.

The stage is a conventional free vortex design with unconventional blades of a special planform. The blade sections of both rotor and stator are NACA 65-sections on circular arc mean lines. Sectional diffusion factors and de Haller numbers are close to their respective limits, especially for the sections next to the rotor and stator hubs.

The rotor is characterised by a forward-swept leading edge with increasing sweep angle towards hub and tip and an unswept trailing edge. The leading edge of the stator blades is forward-swept as before but this time at an almost constant sweep angle between the hub and the two-thirds position of the blade span. The trailing edge is straightened for reducing the previously mentioned aerodynamic loadings.

The study shows that the numerical results are consistent with the experimental outcome. It concludes that the advanced design features show potential aerodynamic and acoustic benefits by sweeping the blade in the described manner. This is particularly the case when comparing to single row designs.

Commentary by Dr. Valentin Fuster
2016;():V001T09A005. doi:10.1115/GT2016-56862.

Market requests for higher performance fans and legal requirements to meet minimum efficiency grades drive industrial fan designers for tunnel and metro applications to study or re-think stall control solutions. This paper, therefore, focuses on how anti-stall rings work and provides a robust and fast numerical methodology to evaluate and improve their efficiency and effectiveness. Here we present a numerical study validated against available experiments on the fitting of an anti-stall ring on an axial fan for tunnel and metro operations. The study synthetizes the effects of the anti-stall ring with an actuator disk coupled to a frozen rotor model for the fan, significantly reducing the computational load required and validating the methodology. Such approach can be implemented to easily derive fan performance with different geometries of the anti-stall ring plenum as well as its fins.

Commentary by Dr. Valentin Fuster
2016;():V001T09A006. doi:10.1115/GT2016-56960.

Tube-axial fans are widely used in industrial applications because of their compactness, simplicity, and low cost. However, the achievable fan pressure rise is generally penalised by the absence of a straightener and diffuser, and the consequent waste of tangential and axial dynamic pressures at the fan outlet. The corresponding fan efficiency drop might not comply with stringent regulations like the European Directive for energy-related products. Thus, operation ranges of high efficiency need to be clearly defined in the preliminary design phase, especially when constraints on maximum size and/or rotational speed are imposed.

This paper proposes analytical formulas and charts to evaluate the efficiency of the tube-axial fan configuration (with or without tail-cone diffuser) when constraints on fan size and/or speed are additional design requirements. The analytical formulas and charts have been validated against experimental data.

On this basis, a preliminary design criterion is suggested for high-efficiency tube-axial fans featuring arbitrary vortex design blades of constant swirl type. The criterion is used to design a 315 mm low-to-medium pressure tube-axial fan that is able to operate at a constant aeraulic efficiency peak of approximately 0.6 for blade positioning angles in the range 20° to 30°.

Topics: Design , Fans
Commentary by Dr. Valentin Fuster
2016;():V001T09A007. doi:10.1115/GT2016-56994.

Both cascade and isolated airfoil methods are considered valid in axial fan blade design, for high (σ≳1) and low (σ≲0.7) solidities respectively. For bladings that feature intermediate solidities the modified isolated approach is commonly employed. This method uses isolated airfoil data, with proper adjustments to take into account multiplane interference effects. Contrarily, the literature does not refer about modifications of the cascade approach to design medium solidity fans. Such method would use cascade data, properly adjusted for the blade sections at lower solidities. Thus, with the aim of comparing these two opposite design approaches (modified cascade versus modified isolated) for medium solidity blades, two free-vortex blading were designed for a 315 mm rotor-only axial fan and experimentally tested. CFD analyses were performed as well to obtain the local flow features. NACA-65 series airfoils were employed, as both cascade and isolated data are available for chord Reynolds numbers typical of axial fans applications. Results highlight the differences between the two approaches. Finally, a mixed approach that employs both isolated and cascade data is suggested as the most accurate one. Moreover, results also show the detrimental effects of the low chord Reynolds numbers on the performance of the blades. This effect should be taken into account in blade design for small-to-medium size machines.

Commentary by Dr. Valentin Fuster
2016;():V001T09A008. doi:10.1115/GT2016-57232.

In the turbo-compressor driving system, working fluid of large capacity and high pressure is stored in the system. Once an accident happened unexpectedly, the driving power of the rotor would be shut off quickly. Owing to the pressure difference between the inlet and the outlet of the compressor, the original downstream fluid might flow back from the volute to the impeller. The backflow might propel the impeller rotating reversely compared with its working status. It would damage the bearing and sealing system. Therefore, it is necessary to avoid the rotor reversal. With the development of the economies of scale, the capacity of the compressor unit becomes larger and larger, the possibility of the rotor reversal accident might be in consideration. The reverse performance of the centrifugal impeller is related to the mass flowrate of the backflow, the backflow pressure, moment of inertia as well as the initial working rotating speed of the rotor. In order to get the quantitative relation of the parameters, a test rig was set up with a centrifugal fan. The parameter was tested on the idle processes with different initial rotating speeds. Compared with the idle process, the reduction processes of the rotating speed driven by the backflow were tested. The processes varied with the pressure and the flowrate of the backflow. In addition, three-dimension simulation on the fan was performed. Corresponding to the experimental data, numerical performance of the fan was got to verify the numerical method. The flow information in the fan with backflow was analyzed. Even the backflow passing through the impeller, there is small flowrate to the volute by the operating fan. Combined with experimental data and the numerical parameters, the theoretical analysis on the reduction process of rotating speed was carried out. The critical values of the initial working rotating speed, the pressure and the flowrate of the backflow were derived to avoid the reverse rotation. Therefore, it can be used for the fan system in the design process to prevent the reverse rotation.

Commentary by Dr. Valentin Fuster
2016;():V001T09A009. doi:10.1115/GT2016-57234.

In recent years there is a growing number of patients suffering from pneumonia, chronic obstructive pulmonary disease or asthma. In these situations, the application of Continuous Positive Airway Pressure (commonly known as CPAP) is indicated by clinicians. It is a noninvasive means of healthcare used widely. As a more advanced technique, the positive airway pressure may follow a time-cycled change between two preset pressure values. This technique is known as BPAP (Bi-level Positive Airway Pressure).

The devices are used mainly during sleep at home. Hence the aeroacoustic requirements are critical. In addition the devices must be portable and compact. Furthermore, the high frequency of pressure change required in BPAP devices poses additional demands on the design. Due to the complexity of the overall design problem, it may be solved efficiently by multi objective optimization.

The pressure head in these devices is generated by radial fans. For the aerodynamic optimization, we utilize a RANS solver. For the aeroacoustic optimization we use the Lattice-Boltzmann Method (LBM). Both are operating on a parametric geometric model of the fan and housing. For the propagation of sound waves into the far-field, we develop algorithmic strategies for using the Ffowcs Williams-Hawkings (FW-H) equation with the LBM. The constrained multi objective optimization is driven by a variant of the NSGA-II algorithm.

We outline the complete optimization procedure for a BPAP device. Our numerical results are compared with physical tests. To analyze the contribution of selected geometric features to the emitted sound pressure, we perform a sensitivity study. The new algorithmic arrangement has shown to drastically cut development costs and time.

Commentary by Dr. Valentin Fuster
2016;():V001T09A010. doi:10.1115/GT2016-57306.

The morphing geometry concept finds interesting applications in load reduction and performance increasing for wings and rotor blades in off-design conditions. Here we report a numerical study on the effect that a passive morphing system (made by an elastic-low stiffness surface) has on the sectional load and flowfield, when it is applied to the trailing edge of an axial fan. We obtain the results extracting the section of the fan blade and test it in the 2D cascade, with and without the elastic device, in different operating conditions. Keeping in mind the two-dimensional approximation, it will be possible to observe how the tested device could reduce the load in off-design and high angle of attack conditions, while the same solution could introduce vibrations in design conditions. All the simulations imply the solution of the fluid-structure interaction between the incompressible, turbulent flow and the elastic structure. This solution is obtained using a finite element based, strongly coupled solver, applied to the periodic 2D domain of the section in the cascade.

Commentary by Dr. Valentin Fuster
2016;():V001T09A011. doi:10.1115/GT2016-57337.

In a traditional automotive cooling system, energy optimization could be achieved by controlling the engine temperature by means of several sensors placed inside the cooling circuit. Nevertheless, in some cases the increasing use of a great number of sensor devices makes the control system too bulky, expensive and not sufficiently robust for the intended application. This paper presents the development of a heavy-duty automotive cooling axial fan with morphing blades activated by Shape Memory Alloy (SMA) strips that work as actuator elements in the polymeric blade structure. The application of smart materials to compact, high-energy density devices as well as the development of modeling and control systems has been of great interest during the last decade. SMAs are frequently combined within monolithic or composite host materials to produce adaptive structures whose properties could be tuned in response to external stimuli.

The blade was designed to achieve the activation of the strips (purposely thermo-mechanically treated) by means of an air stream flow. With the aim of studying the morphing capability of the adaptive structure together with the recovery behavior of the NiTi strips, four different polymeric compounds have been compared in a specifically-designed wind tunnel.

Digital image analysis techniques have been performed to quantitatively analyze the blade deflections and to evaluate the most suitable polymeric matrix for the intended application. As the airstream flow increases in temperature, the strips recover the memorized bent shape, leading to a camber variation. To study the possibility of employing SMA strips as actuator elements, a comparison with common viscous clutch behavior is proposed. The time range actuator response indicates that the SMA strips provide a lower frequency control that fits well with the engine coolant thermal requirement. The experimental results demonstrate the capability of SMA materials to accommodate the lower power actuators in the automotive field. Finally, the blade tip airfoils, reconstructed using a CAD procedure, were used to study the fluid dynamic behavior of the blade tip airfoil. A CFD numerical simulation was carried out in order to highlight the differences in the airfoil performance due to the different shapes of the blade. The analyses showed that the activated blade tip airfoil led to an increase in the lift coefficient according to the stiffness provided by the polymeric compound.

This innovative passive control system results from the selection of (i) the memorized shape of the SMA strips and (ii) the polymeric compound used for the blade structure.

Topics: Fluids , Alloys , Blades , Shapes
Commentary by Dr. Valentin Fuster
2016;():V001T09A012. doi:10.1115/GT2016-57474.

Stall margin improvement is a crucial issue during fan design, selection and installation. In fact, several industrial fans are operated in extreme applications and requested to be highly flexible in order to be able to withstand abrupt changes in operating conditions. This is the case of fans operated under distorted inlet conditions or interacting with other fans or used for reversible operations in tunnel and metro applications.

This paper reports on a systematic experimental study on the effects of the use of casing treatments on the performance of a single stage, reversible, axial fan. Tests focused on the performance of up-stream single treatment and up- and downstream double treatment, respectively designed to prevent the fan to run into stall in forward-only operations and forward- and reverse- operations. The study reports on the assessment of the change in performance, in normal and stalled operations, with an emphasis on the relative axial position between casing treatments and the impeller blades. The analysis demonstrates the achievement of significant improvements in performance in the unstable region for all the tested configurations with marginal losses in the stable counter-part.

Commentary by Dr. Valentin Fuster
2016;():V001T09A013. doi:10.1115/GT2016-57557.

This work investigates the dynamics of rotating stall of a low speed axial fan in presence of fouling on the blades. Rotating stall is an aerodynamic issue of recognized importance in turbomachinery. The combination of rotating stall and presence of particles of dust and dirt from the surrounding environment, may lead to further issues in terms of performance, stall limit and blades life.

In this paper the identification of the rotating stall pattern is carried out using time-resolved sound measurements in the far field region by means of a condenser microphone. The experimental tests are carried out with various geometries of fouling in order to evaluate the system ability to detect acoustically fouling and rotating stall.

The results have been validated against state of the art techniques described in the literature. The acquired signals have been analysed using frequency domain analysis, and time domain analysis using a phase space reconstruction inspired technique. Both of the approaches demonstrate a modification of the stall dynamics in the low speed fan and allow the identification of diverse stall precursors and fouling presence.

Commentary by Dr. Valentin Fuster
2016;():V001T09A014. doi:10.1115/GT2016-57772.

The paper presents a comparative case study in which a free-inlet, free-outlet industrial ventilating fan has been equipped with various inlet geometries. The original short-tapered entry has been replaced by a standardized bellmouth entry, resulting in remarkable noise reduction. The experimentation presented herein is adaptable to industrial onsite diagnostics. The upstream-radiated broadband noise associated with rotating sources has been mapped in a spatially resolved manner, by means of a Phased Array Microphone system and a Rotating Source Identifier beamforming algorithm. The acoustic measurements have been supplemented with aerodynamic measurements on the inlet velocity profile, and with Computational Fluid Dynamics studies. The acoustic data have been processed for enabling their evaluation in association with the aerodynamic operation of the elemental rotor cascades in a two-dimensional approach, and also for their interpretation in relationship to three-dimensional flow phenomena such as tip leakage flow. For this purpose, the acoustic data have been presented in the form of circumferentially-averaged noise profiles along the blade span, as well as noise source imaging maps. The studies reveal the following acoustic benefits of reconfiguring the original short-tapered entry to the bellmouth entry. A peripheral separation zone is characteristic for the short-tapered entry, provoking double-leakage tip clearance flow, being the dominant source of noise at higher frequencies. Such a peripheral separation zone is suppressed by the bellmouth inlet, and thus, the double-leakage flow and the related noise is eliminated. Farther away from the tip, along the dominant portion of the span, the moderation of endwall blockage due to suppressing the peripheral separation zone has led to the reduction of the rotor inlet velocity, thus moderating the noise associated with the suction side boundary layer developing on the blades.

Commentary by Dr. Valentin Fuster
2016;():V001T09A015. doi:10.1115/GT2016-57786.

The acoustical characteristics of fans are an essential criterion of product quality and are continually growing in importance as for example cooling fan modules in the automotive industry have to suffice high comfort requirements. In order to locate dominant acoustic sources and to reduce the noise emission generated by a shrouded fan configuration, numerical simulations are performed. The working approach considers variously modified fan geometries and their evaluation regarding arising vortex flow phenomena and their effect on a decreased sound pressure level (SPL) in consideration of an improvement or the constancy of aerodynamic fan performance. Particular emphasis lies on the analysis of secondary flows in the blade tip region by postprocessing CFD-results. Due to the large number of geometrical modifications investigated and the importance of highly resolved eddy structures, a hybrid approach is chosen by applying the SAS-SST turbulence model in URANS simulations. The SAS (Scale Adaptive Simulation) delivers LES (Large Eddy Simulation) content in unsteady regions of a RANS-simulation and exhibits not nearly the high computational effort needed to perform a full scale LES. An assessment of the actual propagation of noise emission into the far-field is made by performing experimental investigations on the most promising modifications. The acoustic measurements are carried out in a fan test stand in the anechoic chamber of the University of Applied Sciences Duesseldorf. The aerodynamic performance is measured in a fan test rig with an inlet chamber setup in accordance with ISO 5801. The measured acoustical and aerodynamic performance is validated by the industrial partner. The results of the acoustic measurements are in turn utilized to determine indicators of noise radiation in the numerical simulation. Within this work an approach is presented where the analyses of secondary flows in the blade tip region provide the basis for an innovative noise reduction design in a shrouded fan configuration. The new design exhibits a reduced SPL (A-weighted) of approx. 4 dB over the entire operating range while showing no significant deterioration on the aerodynamic performance. The design was registered for patent approval cooperatively by the industrial partner and the University of Applied Sciences Duesseldorf, while minor design parameters are still subject to further improvements. All numerical simulations are performed with Ansys CFX, a commercial solver widely spread in the industry. Methods similar to those shown in this work can be implemented in the design phase of axial fans in order to develop acoustically optimized fan geometries.

Commentary by Dr. Valentin Fuster
2016;():V001T09A016. doi:10.1115/GT2016-57820.

The advancements in fan technology are nowadays animated by two major drivers: the legal requirements that impose minimum fan efficiency grades for fans sold within European Union (and soon US and Asia), and the market request for better air performance and lower sound emissions.

Within HVAC (Heating, Ventilating and Air Conditioning) applications, centrifugal fans with forward curved blades are widely used due to the higher total pressure rise capability and lower acoustic emissions with respect to more efficient backward curved blades. However the continuous rise of minimum fan efficiency grades pushes the manufacturers to develop a new generation of forward curved centrifugal fans, improving previous design. Here the challenge is not only on aerodynamics, but in the overall production process, as squirrel cage fans are characterised by a cost-effective consolidated technology, based on simple blade geometries and easy series manufacturing. For example, the blades usually have circular camber lines, as results of cut cylinders. Thus, once the number of blades and the angle at the leading edge are selected, the chord and the deflection capability are constrained as well.

These concurring aspects led industry to include in the design process new tools, in particular CFD, to analyse the flow features of the current generation of fans in order to understand which phenomena are to be either controlled or exploited to increase efficiency and total pressure rise.

Here we present a numerical investigation on a forward curved blade centrifugal fan for HVAC applications, to highlight the flow features inside the impeller and in the critical region of coupling with the volute. The analysis was carried out with OpenFOAM, an open-source library for CFD. Computations were performed with the frozen rotor approach and validated against available experimental data.

Commentary by Dr. Valentin Fuster


2016;():V001T22A001. doi:10.1115/GT2016-56073.

Gas turbine engines (GTE) of complex cycles are more efficient in comparison to GTE of simple cycle, and they can be recommended for application in a ship propulsion complex. GTE’s representative of complex cycle is GTE with overexpansion turbine and heat regeneration, which has high efficiency, and which can provide heat energy to the consumers of specialized ships in co-generative mode of operation.

GTE with R and OT is adapted for used in different conditions of ships’ operation. It has high power efficiency at partial loads due to the working process control in the engine. There are two applicable methods of getting required engine’s characteristics of GTE with heat regeneration at partial loads.

If more constant production of heat energy is required at partial loads of the main engine, the method of air by-pass to the combustion chamber by recuperator is applied. Hereby, utilized heat is being redistributed towards the gas cooler boiler-utilizer increasing its power.

For the increase of efficiency of GTE with R and OT at partial loads it is used the method of GTE’s characteristics change by application of variable area nozzles (VAN) of power turbine. VAN in GTE with R allows keeping initial temperature of gas in the engine close to nominal value at partial loads. If the efficiency of GTE with R and OT at nominal load is 20 % relatively higher, in this case when power is equal to 50%, the efficiency of GTE with R and OT is 30% higher in relation to Brayton cycle. In GTE with R and OT it is possible to apply two methods of working process control simultaneously in various regimes of ship operation.

Commentary by Dr. Valentin Fuster
2016;():V001T22A002. doi:10.1115/GT2016-56220.

Due to the advantages of great power, small size and light weight, the application of gas turbines in marine power facilities has increased a lot over time. Concerning the factors of off-shore operating environments which include the huge air intake quantity, the massive mist marine circumstance and the precious vessel available capacity, this paper combined experimental study with numerical simulation to investigate the performance of a type of new air intake wave-plate separator which can be applied under high intake velocity conditions. The total pressure drop was measured in a small wind tunnel with the inlet velocities of the separators ranging from 1.0 to 15.0 m/s. The resistance characteristics of high velocity wave-plate separators were simulated under the same velocity range described above. The separation efficiencies of high velocity wave-plate separators were simulated under the inlet velocity of 14.0 m/s, and the liquid diameters were 5μm, 10μm, 15μm and 20μm respectively. By analyzing the results of experiments and simulations, this paper draws the conclusion that high velocity wave-plate separators can keep high separation efficiencies and acceptable total pressure drop under high inlet velocities.

Commentary by Dr. Valentin Fuster
2016;():V001T22A003. doi:10.1115/GT2016-56446.

The MT30 marine gas turbine has been developed specifically for 21st century naval propulsion using modern techniques and methods. Design and development of the MT30 began in 1999 and has since been qualified for naval service following extensive testing. Since then the engine has rapidly been adopted by progressive navies, in both its mechanical and electrical power generation configuration.

The Lockheed Martin Littoral Combat Ship (LCS) is one of a new class of United States Navy (USN) fast combatants which has been at sea for more than six years and is powered by the MT30. A combined MT30-driven generator was selected for the new USN DDG1000 Zumwalt class of destroyer and has also been successfully installed into the Royal Navy’s Queen Elizabeth Class aircraft carrier. Most recently, the MT30 Compact Package has been selected to power the Royal Navy’s Type 26 Global Combat Ship which will be built by BAE Systems.

The MT30 Compact Package has been designed with the aim of powering modern warship programmes, with the result that it is currently the World’s most power dense in-service marine gas turbine. This is an important factor in naval propulsion where delivering a high power output in a compact space is essential.

In addition to the programmes stated above, the MT30 Compact Package was selected for the new Republic of Korea Navy’s (RoKN) frigate programme with a single-GT CODLOG hybrid arrangement consisting of propulsion motors and a Diesel-electric system.

As a result, Rolls Royce was selected by the RoKN to deliver the MT30 Gas Turbine Unit and, from a preliminary Rolls-Royce compact package design, the engine and machinery division of Hyundai Heavy Industry (HHI-EMD) has developed the Compact Package for the New Korea Frigate. The MT30 GT was delivered to the HHI-EMD facility in 2014 with the surrounding Compact Package built at HHI-EMD before onward delivery to Daewoo Shipbuilding and Marine Engineering (DSME) where construction of the first frigate will take place.

This paper provides the rationale for selection of the MT30 Compact Package for the New Korea Frigate Programme and also describes the development of the MT30 Compact Package; aspects of the design process, construction of the Compact Package and the factory acceptance test conducted at the HHI-EMD facility.

Topics: Gas turbines
Commentary by Dr. Valentin Fuster
2016;():V001T22A004. doi:10.1115/GT2016-56724.

When a marine gas turbine is installed on the ship, the alignment among the power turbine, the output shaft and the reduction gearbox needs to be examined and adjusted precisely. After a long period of gas turbine operation, the alignment of the shafting needs be reviewed and adjusted regularly according to relevant regulations as well.

The paper introduces an alignment monitoring device, by which the alignment of the shafting under static and dynamic conditions can be acquired at any time. The device has been verified by real shipboard tests and the on-condition maintenance can be realized. The device can reduce significantly the workload of scheduled maintenance and can be used as a recorder for the responds of gas turbine against various dynamic shocks.

Commentary by Dr. Valentin Fuster
2016;():V001T22A005. doi:10.1115/GT2016-56726.

Gas turbines are widely used as the marine main power system with its higher power density, react quickly, such as LM2500 and MT30. However, it works under design conditions only during running times of 3% to 10%, and it works under part load during most of the time, leading to low efficiency, and it could not achieve full speed or braking at an instant if sudden emergencies happen. Variable geometry turbines can improve this condition by variable angle nozzle (VAN) technology. And, it could enhance engine braking ability, reduce the fuel consumption under part load, improve the aerodynamic performance of engines, enhance accelerating ability of engines, and implement stalling protection to the power turbine.

However, the VAN adjustment needs complicated regulating systems, which makes it difficult to turbine structural design, and leads to increased weight. Besides, there is a performance penalty associated with the vane-end part radial clearance required for the movement of variable vanes.

In order to increase the part load efficiency of an intercooled recuperated gas turbine, the power turbine is converted from fixed to variable geometry. And, in order to reduce the losses caused by the radial clearance both of vane ends while vane turning, spherical ends are introduced to keep the clearance constant at all turning angles, and the baseline clearance is 0.77% of blade span.

In order to determine the effects of VAN on aerodynamic performance of a variable vane, experimental investigations with a variable geometry turbine annular sector cascade have been conducted under five different turning angles (−6°, −3°, 0°, +5° and +10°) and three Mach numbers (0.3Ma, 0.5Ma and 0.6Ma). The parameter distributions were measured at cascade downstream by a five-hole probe and three-axis auto-traversing system, including outlet flow angle, total pressure loss coefficient, energy loss coefficient.

The sector measurement results show that, as the vane turning angle is changed from closed to open, the outlet flow angle are increased under all three test Mach number conditions, which affects the flow mismatching between variable vane and downstream row. And, the total pressure losses is increased with the turning angle changed from design to closed or open, and the total pressure loss increases much more when the vane is closed than when it is open. In addition, vane-end clearances have significantly effects on the flow field. Especially on the hub, the leakage loss is higher, that may be due to the adverse effect of intermediate turbine ducts. Detailed results about these are presented and discussed in the paper.

Commentary by Dr. Valentin Fuster
2016;():V001T22A006. doi:10.1115/GT2016-57214.

Naval ships, as well as commercial ships, are exposed to more risks than before with the development of IR-guided threats. IR Suppress System (IRSS) is used to reduce or eliminate the infrared signatures of exhaust system of the ship. In order to optimize the structure of exhaust ejector and shorten the cycle of research, it is essential to study the successful experiences on exhaust ejectors by scholars abroad. In this article two structure innovations are introduced: an improved structure of nozzle and multi-stage diffuser; A new type of exhaust ejector for marine gas turbine is designed. Through simplifying the ejector model, numerical simulation is applied to forecast the characteristics of both the single-stage ejector and the multistage ejector. Result indicates the effects of structural parameters on the performance parameters. This research result has certain reference value for marine gas turbine exhaust ejector design and performance optimization.

Commentary by Dr. Valentin Fuster
2016;():V001T22A007. doi:10.1115/GT2016-57819.

The Synchro-Self-Shifting (SSS) overrunning clutch is well known, particularly in the Naval Marine field. This paper reviews the clutch operating principle, then outlines some of the service experience since 1962, particularly in naval main propulsion drives beginning with CODOG, CODAG, COGOG and COGAG plant, and then the experience with more recent applications such as combined electric motor propulsion with either gas turbines or diesel engines and hybrid electric plants.

Extra features are then described such as a lock-out control as is usually necessary for turbine applications to permit turbine testing, e.g., when in harbor; also a lock-in control as is essential when the clutch has to transmit power in both directions of rotation. Various clutch mounting arrangements will be presented with respective advantages. The paper concludes with information regarding reliability during many years of service experience.

Commentary by Dr. Valentin Fuster
2016;():V001T22A008. doi:10.1115/GT2016-57991.

The thermal stability of three Ni-base samples was assessed at 1850F (1010°C) and 2000F (1093°C) in ambient air as a function of exposure time ranging from 500 to 2000 hrs. Assessments of thermal stability of the samples were made using weight change, oxidation, microstructural evolution, and post-exposure mechanical properties such as Vickers microhardness and compressive yield stress. The three samples included bare Alloy “A” (9Cr-6Al-1.5Hf), Alloy “A” with an overlay coating, and bare Alloy “B” (12Cr-3Al), were not much different in compositions. At 1850F, oxidation as measured by weight change was insignificant up to 2000 h in all the three samples. At 2000F, however, noticeable weight change occurred, increasing linearly with time all in the three samples. The oxidation penetration from surface to matrix for these samples was more intense when exposed to above 1000 hours, forming various oxides, gamma-prime (γ′) depletion zones, and TCP phases. The size and area fraction of γ′ precipitates were determined as a function of temperature and exposure time. Post-exposure mechanical properties were also assessed through Vickers hardness and compressive yield stress. A maximum change in Vickers hardness was about 10% at both temperatures up to 2000 hrs. The change in compressive yield stress was more pronounced than the change in Vickers hardness as a function of thermal exposure and time.

Commentary by Dr. Valentin Fuster
2016;():V001T22A009. doi:10.1115/GT2016-58135.

With the rapid improvement of equipment manufacturing technology and the ever increasing cost of fuel, engine health management has become one of the most important parts of aeroengine, industrial and marine gas turbine. As an effective technology for improving the engine availability and reducing the maintenance costs, anomaly detection has attracted great attention. In the past decades, different methods including gas path analysis, on-line monitoring or off-line analysis of vibration signal, oil and electrostatic monitoring have been developed. However, considering the complexity of structure and the variability of working environments for engine, many important problems such as the accurate modeling of gas turbine with different environment, the selection of sensors, the optimization of various data-driven approach and the fusion strategy of multi-source information still need to be solved urgently. Besides, although a large number of investigations in this area are reported every year in various journals and conference proceedings, most of them are about aeroengine or industrial gas turbine and limited literature is published about marine gas turbine. Based on this background, this paper attempts to summarize the recent developments in health management of gas turbines. For the increasing requirement of predict-and-prevent maintenance, the typical anomaly detection technologies are analyzed in detail. In addition, according to the application characteristics of marine gas turbine, this paper introduces a brief prospect on the possible challenges of anomaly detection, which may provide beneficial references for the implementing and development of marine gas turbine health management.

Topics: Gas turbines
Commentary by Dr. Valentin Fuster

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