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ItemOptimisation of ducted fans with CFD( 2015) Biava, M. ; Carrión, M. ; Steijl, R. ; Barakos, G.N. ; Stewart,D.This paper presents the performance analysis and design of a ducted propeller for lighter-than-air vehicles. Highfidelity CFD simulations were undertaken on a model with simplified geometry to quantify the effect of the duct, and to assess the influence of the blade twist on the ducted propeller performance. It is shown that the duct is particularly effective for low speed operations, and that the blades with relatively high twist have better performance over a wide range of operating conditions. Design of the optimal twist distribution was then attempted, by coupling the flow solver with a quasi-Newton optimisation method. Flow gradients were provided by a fully implicit adjoint solver for the RANS equations, which accounts for the turbulence model coupling. Results show a 2% reduction in the required power of the optimised ducted propeller. The degree of approximation introduced by the simplified geometry was also quantified, by solving the flow for a more realistic geometry and through comparison with available experimental data.
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ItemEdge-based approach to estimate the drift of a helicopter during flight( 2015) Gatter, A.The Institute of Flight Systems at the German Aerospace Center (DLR) site in Braunschweig Germany has set its goal into making helicopter flying as safe as possible. The new DLR research project "Rettungshubschrauber 2030" addresses the topic of aiding helicopter rescue missions. Research will be conducted to increase the safety of these missions as well as to enable the conduct of missions in circumstances where nowadays a helicopter would not be allowed to operate. One aspect of this research is to increase or maintain the situational awareness of the pilot by processing data from camera images. The presented paper will focus on the field of visual odometry. Most of the publications on this topic use techniques that are only working with satisfying reliability in a very restricted environment, i.e. in good weather conditions. It shall be surveyed, if an edge-based approach for extracting features is a possible alternative or addition to established feature extractors. In the following paper, two algorithms for edge-extraction will be compared: An algorithm that is based on Hough transform and an algorithm that is based on the Douglas-Peucker-Method. They will be tested on their ability to detect a sufficient amount of features in camera images as well as on their computational complexity. Then, their ability to detect the drift of a helicopter will be surveyed on recorded data from flight tests with the Advanced Control Technology/Flying Helicopter Simulator (ACT/FHS) of the DLR. Their performance will be tested on the base of reference data from the ACT/FHS which have been recorded by the use of a highly accurate INS/DGPS system. Finally, a short outlook in form of a first comparison of well established feature extractors and the presented algorithms will be shown on a recorded scene with raindrops covering the lens of the camera.
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ItemUnsteady boundary layer transition on the DSA-9A rotor blade airfoil( 2015) Richter, K. ; Koch, S. ; Goerttler, A. ; Lütke, B. ; Wolf, C.C. ; Benkel, A.The unsteady boundary layer transition on the pitching helicopter main rotor blade airfoil DSA-9A was experimentally investigated by the use of hot film anemometry and unsteady pressure measurements. The unsteady flow characteristics on the upper and the lower side of the airfoil were analyzed for steady test cases and dynamic cases with sinusoidal pitching motion in attached flow conditions at M = 0:30 and Re = 1:8_106. The paper discusses the unsteady transition characteristics in detail and presents the influence of the pitching frequency on the unsteady transition. The results indicate that a large transition hysteresis exists on both sides of the airfoil, and that the hysteresis is much larger than can be explained by the unsteadiness in lift. Significant transition zones exist on the airfoil with sizes of up to 55% chord. The frequency influence is seen in an increase in the hysteresis and in a reduction of the size of the transition zone with increasing frequency.
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ItemDynamic stall simulations with finite-volume based Lattice Boltzmann method( 2015) Guzel, G. ; Yagiz, B. ; Cetiner, A.E.In this study, the capability of the Lattice-Boltzmann Method (LBM) is demonstrated simulating one of the most challenging problems in rotor aerodynamics, i.e. dynamic stall. For this purpose, 2-D simulations with an in-house developed finite-volume based LBM flow solver were performed for a NACA 0015 airfoil that is sinusoidally pitching around its quarter-chord. For the current study, three cases for different mean angles of attack but with the same pitching amplitude and the frequency were considered. Each case corresponds to different flow regimes, i.e. attached flow, light stall, and deep stall. Once the converged solutions were obtained, the computed variations of forces and pitching moment were compared with the measured data and satisfactory results were obtained. Also, the same simulations were repeated with a NS equations based flow solver that is available commercially (Fluent Version 14.5) and the comparison between the two methods revealed that both provides almost equivalent results.
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ItemRotorcraft applications of active gurney flap investigated within the clean-sky project COMROTAG( 2015) Stalewski, W.Computational investigations within the EU 7th FWP Project COMROTAG have been carried out. The main goal of the Project is to develop, implement and validate through computational tests the methodology of simulation of the flow around helicopter rotors with blades equipped with the Active Gurney Flaps (AGF). The AGF is a small, flat tab located at lower surface of the blade, near its trailing edge. The tab is cyclically deployed and retracted perpendicularly to the blade surface. When deployed, the tab deflects air stream behind the blade trailing edge downwards, leading to the lift increase, which is especially important on the retreating blade of the rotor. On the advancing blade, the tab is retracted to minimise rotor torque. The AGF may be potentially used for the active control of flow on helicopter-rotor blades. The paper presents results of already finished stages of the COMROTAG project. The initial stage was focused on development of innovative methodology of computational simulation of flight of the helicopter main rotor with blades equipped with AGF. In general, the methodology is based on solution of the Navier-Stokes Equations via Finite Volume Method. The motion of helicopter rotor blades equipped with dynamically deployed AGF is modelled using Deforming Mesh and Sliding Mesh techniques. First stage of validation of the developed methodology concerned quasi-2D investigations of rotor-blade segment equipped with the AGF. The computational results have been compared with experimental data obtained in wind-tunnel tests.
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ItemNumerical simulation of the ERICA tiltrotor using the HMB2 solver( 2015) Garcia, A.J. ; Barakos, G.N.Numerical simulations of the Enhanced Rotorcraft Innovative Concept Achievement tiltrotor have been performed using the Helicopter Multi-Block CFD solver. Comparisons with experimental data obtained in the German-Dutch Large Low-speed Facility and the S1MA wind tunnel are also presented. For this work, an aeroplane mode configuration, referred to as AC1, of the tiltrotor is considered. It is characterised by low speed and relatively high angle of attack of the aircraft. The Helicopter Multi-Block CFD method predicted the surface pressure coefficient at cross sections on the fuselage, nacelle, fixed wing, and tiltable wing of the tiltrotor accurately. Good overall agreement is obtained between CFD and wind tunnel data. In addition to fully resolved blades, computations with actuator disk models were also performed.
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ItemCFD-simulation of three-dimensional dynamic stall on a rotor with cyclic pitch control( 2015) Letzgus, J. ; Keßler, M. ; Krämer, E.Computational fluid dynamics (CFD) simulations of a two-bladed Mach-scaled rotor (R = 0.65 m, Matip = 0.6, Retip _ 1 × 106) with 1/rev cyclic pitch control encountering three-dimensional dynamic stall are presented. The block-structured flow solver FLOWer is used along with the Menter SST turbulence model and a fifth-order spatial CRWENO scheme. A grid and time step dependency study shows the need of high resolution in space and time to properly resolve all stages of dynamic stall. The onset of flow separation in the outboard region of the rotor blade during its upstroke is found to be shock-induced. Flow pattern visualizations reveal the following evolution of a discrete -shaped vortex. Quickly thereafter a partially spanwise vortex occurs, which is bent towards the leading edge near the mid-span. An interaction with the blade tip vortex is noticed as a limitation of outboard-faced separation spreading. Only a small range of outboard radial sections could be found at which the flow pattern resembles two-dimensional dynamic stall. The superposition of an axial flow weakens the dynamic stall event and slightly changes the vortex pattern.
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ItemStructural demonstration of an active back-flow flap under wind tunnel conditions( 2015) Opitz, S. ; Gardner, A.D. ; Merz, C.B. ; Wolf, C.C.The paper presents a detailed concept for influencing dynamic stall with a surface integrated active back-flow flap on the upper side of an airfoil. The development of the flap from a basic concept to the final wind tunnel experiment is described. Special attention is paid on the selection of flap size and position, the structural concept, the actuation mechanism and the instrumentation. Further, the manufacturing procedure developed to produce a retrofit solution for an existing wind tunnel model is illustrated. The paper closes with first functional tests of actuation mechanism and instrumentation in the wind tunnel. Finally the results from wind tunnel experiments are used to validate the predicted reduction of the pitching moment peak.
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ItemPropeller analysis using a hybrid Navier-Stokes/Free-wake method( 2015) Min, B.-Y. ; Brian E. Wake, B.E. ; Bowles, P.O. ; Matalanis, C.G. ; Moffitt, B.The test bench providing application of experimental technology includes operating helicopter main rotor and three blade models.
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ItemLead-lag dynamics of a rotor with stick-slip nonlinearity( 2015) Andersch, P. ; Hajek, M.An assessment of a frictional helicopter blade attachment with regard to its influence on the dynamics of the in-plane blade motion is presented. For detailed system analysis, a 3D finite element model with contact modeling has been created and validated using preexisting test data. Using the insights gained from 3D finite element simulations, a reduced order model suitable for dynamical simulations is derived and parameterized. Eventually, the reduced order model for the blade attachment is inserted into a hingeless dynamical rotor model with elastic coupling. The influence of the blade attachment on lead-lag dynamic behavior is investigated using the developed model and multiple configurations are compared to each other to identify the effects of the nonlinear blade attachment model on lead-lag frequency and damping ratio as well as elastic coupling prevalent in the hingeless rotor configuration
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ItemTime-and-spatially adapting simulations for efficient dynamic stall predictions( 2015) Smith, M.J. ; Jain, R. ; Grubb, A. ; Jacobson, K.The ability to accurately and efficiently predict the occurrence and severity of dynamic stall remains a major roadblock in the design and analysis of conventional rotors as well as new concepts for future vertical lift. Several approaches to reduce the cost of these dynamic stall simulations for airfoils and finite wings are investigated. Temporal error controllers, variable time step sizes, and feature-based near-body mesh adaptation are evaluated for their ability to more cost-effectively predict dynamic stall on three different configurations. A fourth-order temporal controller has been observed to provide a balanced cost-accuracy ratio, as a maximum of three to four orders of magnitude convergence of the Newton subiterations is obtained during much of the dynamic stall cycle. Larger times steps can be applied, in particular during the attached upstroke portion of the dynamic stall cycle with fourth-order temporal convergence. Mesh reductions via a feature-based two-level adaptation provided a 50% reduction in computational costs with comparable accuracy to a fixed, refined mesh size. Additional refinements may be warranted just after the dynamic stall onset to capture the complex flow field.
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ItemTechnological tests regarding the qualification of the REACh compliant surface treatment processes for helicopter dynamic systems( 2015) Güzel, A. ; Heredia, S. ; Besson, J-M.REACh regulation imposes to European industrial companies to apply viable alternatives for the components which contain hexavalent chromium from 2017. Hexavalent chromium is mainly used in surface treatment operations, which affect the strength relevant properties of the helicopter components, such as the fatigue behaviour as well as the resistance against wear, fretting and corrosion. The components of the helicopter's dynamic system treated by alternative surface treatment processes have to be certified after the qualificationof the REACh compliant processes. This paper presents a technological test program used for determining the performances of the proposed REACh compliant surface treatment processes compared to the current ones. One key issue for the dynamic system components is the contact between two components for which new surface treatments will be applied. In order to determine the contact pairs to be tested, a complete screening of the critical components of the dynamic system has been performed, whose failure is catastrophic (e.g. rotors, gearboxes, etc.), and representative technological samples have been designed. This technological test program includes the fretting wear tests for different plane to plane contacts, fretting fatigue tests in cylinder to cylinder contacts, corrosion tests inside bolted connections and the combined wear-corrosion tests. The principle is to compare each time the current design with the REACh substitutes. Depending on these tests results, some complement can be done in case of differences observed or some additional specific tests can be required based on the service experience coming from the development and/or major incidents. The results of the technological test provide not only a performance check required for the certification of the REACh compliant products, but also an extensive experimental knowledge on the fatigue, wear, corrosion and fretting resistance of the materials used in the aerospace industry.
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ItemWake characteristics of large-scale wind turbines( 2015) Wang, Y. ; Leble, V. ; White, M. ; Barakos, G.N.The next generation of large-scale wind turbines will exceed 10 MW of rated power and will reach rotor diameters of about 200 m. Their rotor aerodynamics are also extreme with Reynolds numbers that reach 40 million. The wakes generated by these wind turbines cover a very large area downstream of their installation positions which increases the possibility that the wake vortices generated by these large wind turbines may affect passing-by flying vehicles. In this paper a CFD study of a large wind turbine was carried out to predict the power curves and aerodynamic loads on the rotor blades. Flow control devices of active leading and trailing edge flaps were also considered in the CFD study and the effects of flaps were investigated. The near wake flow was captured in the CFD study and the flaps add more complexity to the wake flow. To study the potential wind turbine wake encounters by aircraft, engineering wake models were developed to predict the wind turbine far wakes. The wake induced velocity fields were integrated into an aircraft flight dynamics model to simulate wind turbine wake encounter scenarios, designed for a light aircraft approaching an airport, where a wind turbine is installed. The severity of the wind turbine wake encounter was analysed using off-line flight simulations. The off-line simulation results indicated that the wake encounter severity was highly dependent on the ways that the wake vortex circulation and the core size were calculated, which suggested that field measurements of large wind turbines wake flow are needed to verify the modelling and CFD results.
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ItemMarker layer technology: advanced crack detection in elastomeric bearings( 2015) Halladay, J.R.Helicopter elastomeric bearings are typically replaced "on condition", that is, once elastomer fatigue degradation has reached a prescribed depth, the bearings are removed from service. To simplify inspection and enable proactive maintenance, LORD Corporation has developed marker layer technology using a contrasting color elastomer to visually indicate that a bearing is approaching the end of its service life.
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ItemBlue Pulse™: Active rotor control by trailing edge flaps at AIRBUS HELICOPTERS( 2015) Dieterich, O. ; Rabourdin, A. ; Maurice, J.-B. ; Konstanzer, P.Ten years ago, AIRBUS HELICOPTERS has opened a new page in active rotor control by performing the first flight of a main rotor system equipped with piezo-driven trailing edge flaps, In 2005, the experimental test bed consisted of a modified BK117 composite airframe and the experimental ADASYS main rotor system labelled according to the underlying research project. In the meantime, the experimental system has been continuously improved in order to close the gap to potential serial solutions in view of airframe, main rotor, actuation and mechatronics. Until today, AIRBUS HELICOPTERS is the only manufacturer in the world performing full scale flight tests with such kind of systems. The related technology is branded by AIRBUS HELICOPTERS as Blue PulseTM. This paper provides a comprehensive overview on active rotor activities at AIRBUS HELICOPTERS concerning piezo-driven trailing edge flaps. Main topics are the presentation of flight test results and underlying flight physics for major secondary control tasks. Future development challenges are identified to be less on realization by design and proof-of-demonstration for both hardware and control algorithms but more on maturing the underlying concepts in view of potential customer applications.
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ItemA simple analytical fuselage-induced velocity model for comprehensive rotor codes( 2015) Wall, B.G. van der ; Wentrup, M. ; Rajagopalan, G. ; Jung, S.N.A simple analytical model to account for fuselage-induced velocities at rotor blade elements and at rotor wake nodes is described. The method is applied to three different fuselage configurations. Results obtained with a comprehensive rotor code show the fuselage effect on rotor trim controls, comparing the isolated rotor with inclusion of the fuselage for the same trim. This is compared to a simple analytical estimate of the fuselage effect using blade element/momentum theory. It is found that in forward flight the lateral control is mainly affected by fuselage effects. Rotor thrust can be varied by the presence of the fuselage, depending on its angle of attack, and the fuselage influence generally increases with flight speed.
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ItemIn-flight tuning system for the CH-53G helicopter( 2015) Arnold, U.T.P. ; Fuerst, D.Rotor Track and Balance (RT&B) adjustments manually applied to the rotor on the ground are burdensome, error-prone and time-consuming. Moreover, as often as the rotor properties change for any reason the RT&B procedure must be repeated. Today, the search of an acceptable setting usually requires a sequence of several dedicated check flights in different RT&B flight conditions to verify the effect of those manual changes. Due to the flight-condition-dependent effect of blade dissimilarities and due to the limited number of locations where compensational changes can be applied, the suppression of the 1/rev unbalances can never be perfect across all flight regimes. Therefore, a system designed to adjust the RT&B setting during the flight can significantly improve the over-all vibratory condition of the rotorcraft. In addition, such system considerably reduces the time for RT&B still required today for the repetitive check flights and the manual adjustments. ZF Luftfahrttechnik had been contracted to demonstrate the feasibility of this approach within a dedicated flight test campaign. Core components of the In-Flight Tuning (IFT) system are electrical Smart Pitch Rods (SPR™), which replace the rigid pitch links of each blade. The SPR™s change their length upon digital commands received from a dedicated control computer. Such close-to-production IFT system has been installed onto a CH-53G testbed of the German Armed Forces and was flight tested at the German Military Flight Test Center in Manching. The tests were designed to prove the concept in a realistic environment and have successfully proven autonomous IFT operation based upon adaptive closed loop algorithms. The benefits that were demonstrated during this campaign comprise (1) reduced vibrations compared to the reference case with fixed pitch link settings throughout all flight regimes, (2) compensation of degrading RT&B condition over time, (3) automatic reconfiguration after sudden changes of blade properties, and (4) reduced RT&B effort after initial rotor reassembly, blade exchange and/or control system rigging. It could also be shown that the IFT algorithm was convergent and robust throughout all maneuvers and never lost track even during the most agile flight condition changes.
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ItemA new semi-empirical damage tolerance and fatigue evaluation approach for composite rotorcraft airframe structures( 2015) Engleder, A. ; Francescatti, D. ; Arelt, R. ; Burger, U.A new semi-empirical methodology to perform a damage tolerance and fatigue evaluation for composite airframe Principal Structural Elements (PSE) is proposed, to comply to new CS/FAR requirement §29.573. "Damage tolerance and fatigue evaluation of composite rotorcraft structures". This methodology consists out of five different steps. 1st , identification of airframe PSE based on the consequence of their failure, e.g. by a FMECA. 2nd , individual threat assessment for each PSE, based on in-service experience, in-house and from others. 3rd, determination of detectability thresholds for individual PSE by performing impact tests on specific coupons to derive Barely Visible Impact Damages (BVID)/ Clearly Visible Impact Damage (CVID) detectability thresholds and CVID/Obvious damage detectability thresholds. 4th, no-growth demonstration of damages on impacted coupons by applying repeated loads with constant amplitude for a certain number of load cycles which cover one design service goal (DSG/life). Those test results are then used to derive allowables in terms of strain limits for sizing of PSE. 5th, structural full scale component tests (static and fatigue) of PSE with BVID/CVID and other typical in-service damages and repairs, to verify Design Ultimate Load (DUL) capability, no-growth behavior of damages, suitability of repairs and residual strength capability, which is minimum Design Limit Load (DLL).
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ItemANCORA: Anotec-Comoti rotorcraft acoustics initiative for preliminary acoustic flight tests for the tuning of simplified rotorcraft noise models( 2015) Van Oosten, N. ; Dragasanu, L. ; Cenedese, F.One of the objectives of Clean Sky GRC5 is to implement a tool for the minimisation of noise impact on the ground, capable of being executed on-board "on-the-fly", providing flight directives to the Flight Management System of the helicopter. The semi-empirical model to be used for this purpose requires information to be derived from experimental data. For this purpose noise measurements have to be made simultaneously on the exterior of the helicopter (i.e. close to the noise source) and on the ground. The main objective of the ANCORA project was to develop the measurement systems and methodologies required to derive the transfer functions between on-board and ground-based microphones and validate them with flight tests.
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ItemRobust aerodynamic optimization strategies for rotor blade morphing airfoils( 2015) Fusi, F. ; Congedo, P.M. ; Guardone, A. ; Quaranta, G.An aerodynamic optimization method is developed to define robust shapes for morphing airfoils for helicopter blades. The morphing strategy consists of a conformable camber airfoil which changes over the period of rotation of the blade to cope with the variable ow conditions encountered in forward flight. A robust or uncertainty-based approach is used to compute a reliable morphing airfoil, providing a low variance with respect to uncertainty affecting the operating conditions. In order to assess the effectiveness of the robust method, several optimization problems are performed, from a classical two-point drag minimization with lift coefficient constraint to a robust morphing camber optimization. The results of the optimization problems are compared and discussed to highlight the features of the robust approach.