To see the other types of publications on this topic, follow the link: Rotors (Helicopters) - Models.
Dissertations / Theses on the topic 'Rotors (Helicopters) - Models'
Create a spot-on reference in APA, MLA, Chicago, Harvard, and other styles
Consult the top 46 dissertations / theses for your research on the topic 'Rotors (Helicopters) - Models.'
Next to every source in the list of references, there is an 'Add to bibliography' button. Press on it, and we will generate automatically the bibliographic reference to the chosen work in the citation style you need: APA, MLA, Harvard, Chicago, Vancouver, etc.
You can also download the full text of the academic publication as pdf and read online its abstract whenever available in the metadata.
Browse dissertations / theses on a wide variety of disciplines and organise your bibliography correctly.
1
Fulton, Mark V. "Stability of elastically tailored rotor blades." Diss., Georgia Institute of Technology, 1992. http://hdl.handle.net/1853/12248.
Full text
2
Usta, Ebru. "Application of a symmetric total variation diminishing scheme to aerodynamics of rotors." Diss., Georgia Institute of Technology, 2002. http://hdl.handle.net/1853/13018.
Full text
3
Wang, Gang. "Study of a low-dispersion finite volume scheme in rotocraft noise prediction." Diss., Georgia Institute of Technology, 2002. http://hdl.handle.net/1853/12395.
Full text
4
Zhao, Jinggen. "Dynamic Wake Distortion Model for Helicopter Maneuvering Flight." Diss., Georgia Institute of Technology, 2005. http://hdl.handle.net/1853/7103.
Full textAbstract:
A new rotor dynamic wake distortion model, which can be used to account for the rotor transient wake distortion effect on inflow across the rotor disk during helicopter maneuvering and transitional flight in both hover and forward flight conditions, is developed. The dynamic growths of the induced inflow perturbation across the rotor disk during different transient maneuvers, such as a step pitch or roll rate, a step climb rate and a step change of advance ratio are investigated by using a dynamic vortex tube analysis. Based on the vortex tube results, a rotor dynamic wake distortion model, which is expressed in terms of a set of ordinary differential equations, with rotor longitudinal and lateral wake curvatures, wake skew and wake spacing as states, is developed. Also, both the Pitt-Peters dynamic inflow model and the Peters-He finite state inflow model for axial or forward flight are augmented to account for rotor dynamic wake distortion effect during helicopter maneuvering flight. To model the aerodynamic interaction among main rotor, tail rotor and empennage caused by rotor wake curvature effect during helicopter maneuvering flight, a reduced order model based on a vortex tube analysis is developed.
Both the augmented Pitt-Peters dynamic inflow model and the augmented Peters-He finite state inflow model, combined with the developed dynamic wake distortion model, together with the interaction model are implemented in a generic helicopter simulation program of UH-60 Black Hawk helicopter and the simulated vehicle control responses in both time domain and frequency domain are compared with flight test data of a UH-60 Black Hawk helicopter in both hover and low speed forward flight conditions.
5
Benjanirat, Sarun. "Computational studies of the horizontal axis wind turbines in high wind speed condition using advanced turbulence models." Diss., Available online, Georgia Institute of Technology, 2006, 2006. http://etd.gatech.edu/theses/available/etd-08222006-145334/.
Full textAbstract:
Thesis (Ph. D.)--Aerospace Engineering, Georgia Institute of Technology, 2007.Samual V. Shelton, Committee Member ; P.K. Yeung, Committee Member ; Lakshmi N. Sankar, Committee Chair ; Stephen Ruffin, Committee Member ; Marilyn Smith, Committee Member.
6
Patrick-Aldaco, Romano. "A Model Based Framework for Fault Diagnosis and Prognosis of Dynamical Systems with an Application to Helicopter Transmissions." Diss., Georgia Institute of Technology, 2007. http://hdl.handle.net/1853/16266.
Full textAbstract:
The thesis presents a framework for integrating models, simulation, and experimental data to diagnose incipient failure modes and prognosticate the remaining useful life of critical components, with an application to the main transmission of a helicopter. Although the helicopter example is used to illustrate the methodology presented, by appropriately adapting modules, the architecture can be applied to a variety of similar engineering systems. Models of the kind referenced are commonly referred to in the literature as physical or physics-based models. Such models utilize a mathematical description of some of the natural laws that govern system behaviors.
The methodology presented considers separately the aspects of diagnosis and prognosis of engineering systems, but a similar generic framework is proposed for both. The methodology is tested and validated through comparison of results to data from experiments carried out on helicopters in operation and a test cell employing a prototypical helicopter gearbox. Two kinds of experiments have been used. The first one retrieved vibration data from several healthy and faulted aircraft transmissions in operation. The second is a seeded-fault damage-progression test providing gearbox vibration data and ground truth data of increasing crack lengths. For both kinds of experiments, vibration data were collected through a number of accelerometers mounted on the frame of the transmission gearbox. The applied architecture consists of modules with such key elements as the modeling of vibration signatures, extraction of descriptive vibratory features, finite element analysis of a gearbox component, and characterization of fracture progression.
Contributions of the thesis include: (1) generic model-based fault diagnosis and failure prognosis methodologies, readily applicable to a dynamic large-scale mechanical system; (2) the characterization of the vibration signals of a class of complex rotary systems through model-based techniques; (3) a reverse engineering approach for fault identification using simulated vibration data; (4) the utilization of models of a faulted planetary gear transmission to classify descriptive system parameters either as fault-sensitive or fault-insensitive; and (5) guidelines for the integration of the model-based diagnosis and prognosis architectures into prognostic algorithms aimed at determining the remaining useful life of failing components.
7
Rauchenstein, Werner J. "A 3D Theodorsen-based rotor blade flutter model using normal modes." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2002. http://library.nps.navy.mil/uhtbin/hyperion-image/02sep%5FRauchenstein.pdf.
Full textAbstract:
Thesis (M.S. in Aeronautical Engineering)--Naval Postgraduate School, September 2002.Thesis advisor(s): E. Roberts Wood, Mark A. Couch. Includes bibliographical references (p. 55-56). Also available online.
8
Sturisky, Selwyn H. "A linear system identification and validaton of an AH-64 apache aeroelastic simulation model." Diss., Georgia Institute of Technology, 1993. http://hdl.handle.net/1853/13402.
Full text
9
Su, Ay. "Application of a state-space wake model to elastic blade flapping in hover." Diss., Georgia Institute of Technology, 1989. http://hdl.handle.net/1853/11965.
Full text
10
Bitzer, Michael. "Identification of an improved body aerodynamics model for the BO 105." Thesis, Georgia Institute of Technology, 1989. http://hdl.handle.net/1853/13832.
Full text
11
Stettner, Martin. "Application of a state-space wake model to a servo flap controlled rotor in hover." Thesis, Georgia Institute of Technology, 1990. http://hdl.handle.net/1853/20202.
Full text
12
Xin, Hong. "Development and validation of a generalized ground effect model for lifting rotors." Diss., Georgia Institute of Technology, 1999. http://hdl.handle.net/1853/11880.
Full text
13
Kaladi, Vasudevan M. "Unsteady compressible lifting surface analysis for rotary wings using velocity potential modes." Diss., Georgia Institute of Technology, 1990. http://hdl.handle.net/1853/12524.
Full text
14
Brand, Albert G. "An experimental investigation of the interaction between a model rotor and airframe in forward flight." Diss., Georgia Institute of Technology, 1989. http://hdl.handle.net/1853/12433.
Full text
15
Thompson, Thomas L. "Velocity measurements near the blade tip and in the tip vortex core of a hovering model rotor." Diss., Georgia Institute of Technology, 1986. http://hdl.handle.net/1853/13003.
Full text
16
Chouchane, Mnaouar. "Application of a dynamic stall model to rotor trim and aeroelastic response." Diss., Georgia Institute of Technology, 1989. http://hdl.handle.net/1853/12368.
Full text
17
Stumpf, Walter Martin. "An integrated finite-state model for rotor deformation, nonlinear airloads, inflow and trim." Diss., Georgia Institute of Technology, 1992. http://hdl.handle.net/1853/13341.
Full text
18
Russell, Gregory T. "Development of an analytical model for pitch link loads of bearingless main rotors." Related electronic resource: Current Research at SU : database of SU dissertations, recent titles available full text, 2007. http://wwwlib.umi.com/cr/syr/main.
Full text
19
Steyn, J. (Johannes). "Design, manufacture and test of a bearingless rotor hub for the 24% scale model of the Rooivalk attack helicopter." Thesis, Stellenbosch : Stellenbosch University, 2000. http://hdl.handle.net/10019.1/51676.
Full textAbstract:
Thesis (MEng) -- Stellenbosch University, 2000.Some digitised pages may appear illegible due to the condition of the original hard copy.
ENGLISH ABSTRACT: This document contains the work done on the bearingless rotor hub for the 24% scale model of the Rooivalk Attack Helicopter situated at the CSIR in Pretoria. This work forms part of the MSc Ing degree of Johannes Steyn. This work was deemed necessary because of a movement away from the fully articulated rotor to one of hingeless and more recently bearingless rotors. The main emphasis of this thesis is to be a technology demonstrator more than the design of a fully working bearingless rotor hub. With this in mind the final design in this report is not an optimal one, but the procedures and methodology in getting to a design are laid out in this document. To verify the design, tests were identified and created. The procedures for these tests are also included in this document. For the fatigue test a test bench had to be designed and built. This document also includes the design of this test bench
AFRIKAANSE OPSOMMING: Die dokument lewer verslag van die aktiwiteite vir die ontwerp van ‘n laerlose rotor van die 24% skaal model van die Rooivalk Helikopter, gelee by die WNNR in Pretoria. Die werk gedoen vorm deel van die MSc Ing graad van Johannes Steyn. Die werk is nodig geag omdat daar ‘n tendens is om weg te beweeg van die volledig geartikuleerde rotor na die van ‘n skanierlose en meer huidig ‘n laerlose rotor. Die hoof klem van die tesis is om as tegnologie demonstrator op te tree, eerder as die daarstel van ‘n werkende laerlose rotor. Na aanleiding van bogenoemde stelling kan die finale ontwerp nie as optimaal beskou word nie. Die prosedures en metodiek wat gevolg is om die ontwerp te kry word uitgele in die dokument. Om die ontwerp te verifieer is toetse gei'dentifiseer. Die prosedures vir elk van die toetse word ook in die dokument ingesluit. Vir die uitputtingstoetse moes ‘n spesiale toetsbank ontwerp en gebou word. Die ontwerp van hierdie toetsbank is ook in die dokument.
20
Sotiriou, C. P. "An experimental and a theoretical investigation of rotor pitch damping using a model rotor." Thesis, University of Southampton, 1990. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.277512.
Full text
21
Ku, Jieun. "A Hybrid Optimization Scheme for Helicopters with Composite Rotor Blades." Diss., Georgia Institute of Technology, 2007. http://hdl.handle.net/1853/16268.
Full textAbstract:
Rotorcraft optimization is a challenging problem due to its conflicting requirements among many disciplines and highly coupled design variables affecting the overall design. Also, the design
process for a composite rotor blade is often ambiguous because of its design space. Furthermore, analytical tools do not produce
acceptable results compared with flight test when it comes to aerodynamics and aeroelasticity unless realistic models are used, which leads to excessive computer time per iteration.
To comply these requirements, computationally efficient yet realistic tools for rotorcraft analysis, such as VABS and DYMORE were used as analysis tools. These tools decompose a three-dimensional problem into a two-dimensional cross-sectional and a one-dimensional beam analysis. Also, to eliminate the human interaction between iterations, a previously VABS-ANSYS macro was
modified and automated. The automated tool shortened the computer time needed to generate the VABS input file for each analysis from
hours to seconds. MATLAB was used as the wrapper tool to integrate VABS, DYMORE and the VABS-ANSYS macro into the methodology. This methodology uses Genetic Algorithm and gradient-based methods as
optimization schemes. The baseline model is the rotor system of generic Georgia Tech Helicopter (GTH), which is a three-bladed, soft-in-plane, bearingless rotor system. The resulting methodology is a two-level optimization, global and local. Previous studies showed that when stiffnesses are used as design variables in
optimization, these values act as if they are independent and produce design requirements that cannot be achieved by local-level optimization. To force design variables at the global level to stay within the feasible design space of the local level, a surrogate model was adapted into the methodology. For the surrogate model,
different ``design of experiments" (DOE) methods were tested to find the most computationally efficient DOE method. The response surface
method (RSM) and Kriging were tested for the optimization problem.
The results show that using the surrogate model speeds up the optimization process and the Kriging model shows superior performance over RSM models. As a result, the global-level optimizer
produces requirements that the local optimizer can achieve.
22
Hanson, Berenike. "Investigation of a non-uniform helicopter rotor downwash model." Thesis, Linköping University, Department of Electrical Engineering, 2008. http://urn.kb.se/resolve?urn=urn:nbn:se:liu:diva-17193.
Full textAbstract:
This master thesis was carried out at the Department of Aerodynamics and Flight Mechanics at Saab Aerosystems, Linköping, Sweden. It makes up the author’s final work prior to graduation in the field Applied Physics and Electrical Engineering at the Department of Electrical Engineering at The Linköping Institute of Technology (LiTH), Linköping, Sweden.
The objective of the paper was to study a non-uniform helicopter rotor downwash model in forward flight for the unmanned helicopter Skeldar, which is under development at Saab. The main task was to compare the mentioned model with today’s uniform downwash model in order to find differences and similarities. This was done both from a modeling and a controlling perspective. To start with, an introduction is given which is followed by a helicopter theory chapter. The following three chapters deal with the theory of induced velocity, the helicopter model and the Linear Quadratic Regulator (LQR). Finally, the results are presented and discussed.
The downwash models were derived using Momentum Theory (MT) and Blade Element Theory (BET). These two theories were combined in order to find a connection between the induced velocity and the rotor thrust coefficient. The non-uniform downwash model that was studied is proposed by Pitt & Peters and describes a linear variation of the induced velocity in the longitudinal plane.
For the control, a LQ-regulator was chosen since it is easily implemented in MATLAB and it stabilizes the plant model by feedback and consequently creates a robust system. Before the controller could be implemented, the models had to be reduced and the states had to be divided into longitudinal and lateral ones.
The comparison between the open systems clearly shows that differences in the inflow models propagate to all states and consequently the helicopter behaves differently in all planes. Great discrepancies are apparent for the angular velocities p and q. For Pitt & Peters’ model those states are believed to be strongly affected by the system’s positive real pole, causing a rather unstable behavior. When the systems were closed by feedback, the differences were reduced dramatically. Pitt & Peters’ model resulted in greater overshoots than the uniform model, but the overall behavior of all states was rather similar for the two models.
It is concluded, that the adaption of Pitt & Peters’ inflow model does not make any substantial difference when a controller is implemented. The differences between the open systems, however, are reason enough to question Pitt & Peters’ model. In order to evaluate the non-uniform model properly, it has to be compared to suitable flight data which is still lacking up to this date.
23
Wing, Eliya. "Numerical simulation of ice accretion on 3-D rotor blades." Thesis, Georgia Institute of Technology, 2014. http://hdl.handle.net/1853/51833.
Full textAbstract:
Rotorcraft vehicles are highly sensitive to ice accretion. When ice forms on helicopter rotor blades, performance degradation ensues due to a loss of lift and rise in drag. The presence of ice increases torque, power required, and leads to rotor vibrations. Due to these undesirable changes in the vehicle's performance, the FAA requires intensive certification to determine the helicopter’s airworthiness in icing conditions. Since flight tests and icing tunnel tests are very expensive and cannot simulate all conditions required for certification, it is becoming necessary to use computational solvers to model ice growth and subsequent performance degradation. Currently, most solvers use the strip theory approach for 3D shapes. However, rotor blades can experience significant span-wise flow from separation or centrifugal forces. The goal of this work is to investigate the influence of span-wise flow on ice accretion. The classical strip theory approach is compared to a curved surface streamline based approach to assess the relative differences in ice formation.
24
Spathopoulos, Vassilios McInnes. "The assessment of a rotorcraft simulation model in autorotation by means of flight testing a light gyroplane." Thesis, Connect to e-thesis, 2001. http://theses.gla.ac.uk/797/.
Full textAbstract:
Thesis (Ph.D.) - University of Glasgow, 2001.Ph.D. thesis submitted to the Department of Aerospace Engineering, University of Glasgow, 2001. Includes bibliographical references. Print version also available.
25
Ellenrieder, Thomas Jochen. "Investigation of the dynamic wake of a model rotor." Thesis, University of Bristol, 1995. http://hdl.handle.net/1983/13c5bb18-5952-41ae-ba03-08d8d2040d12.
Full textAbstract:
In this study the dynamic induced velocity field of a model helicopter rotor - excited in collective and cyclic pitch at frequencies extending to 1.5 times the nominal shaft speed - is investigated using mainly hot-wire and laser Doppler anemometry. The dynamic induced velocities are found to vary significantly with radial station and frequency. For cyclic excitations, azimuthal variations are also observed. The results point to the dynamic induced flow being influenced by the distribution of shed vorticity in the wake and cannot be explained using simple momentum theory. Vertical variations of the measured inflow response are also observed, with phase changes possibly partly due to transmission type delays. At frequencies above shaft speed a change in character of the induced flow is seen and around shaft speed an increase in the general level of turbulence is found. The available data on the dynamic induced velocity field of a rotor under controlled excitation, are substantially extended. The measured induced flow response was compared to that predicted using the Pitt and Peters dynamic inflow model. In the 'collective' case good agreement was found, suggesting that the primary inflow model parameters such as the inflow gain and apparent mass are correct with some evidence that a higher order inflow representation might be desirable. A novel method is used to infer the aerodynamic hub loading, which could not be directly measured, from the blade flapping data. This is used to isolate the inflow response using the Pitt and Peters dynamic inflow model and the results are compared with experimental measurements. The method shows the Pitt and Peters dynamic inflow representation to be adequate in the 'collective' case. In the 'cyclic' case, the inferred hub loads were very sensitive to the blade model and hence conclusions for this case are limited. A literature survey and review of the Pitt and Peters dynamic inflow model are also given
26
Olcer, Fahri Ersel. "Linear time invariant models for integrated flight and rotor control." Diss., Georgia Institute of Technology, 2011. http://hdl.handle.net/1853/44921.
Full textAbstract:
Formulation of linear time invariant (LTI) models of a nonlinear system about a periodic equilibrium using the harmonic domain representation of LTI model states has been studied in the literature. This thesis presents an alternative method and a computationally efficient scheme for implementation of the developed method for extraction of linear time invariant (LTI) models from a helicopter nonlinear model in forward flight. The fidelity of the extracted LTI models is evaluated using response comparisons between the extracted LTI models and the nonlinear model in both time and frequency domains. Moreover, the fidelity of stability properties is studied through the eigenvalue and eigenvector comparisons between LTI and LTP models by making use of the Floquet Transition Matrix. For time domain evaluations, individual blade control (IBC) and On-Blade Control (OBC) inputs that have been tried in the literature for vibration and noise control studies are used. For frequency domain evaluations, frequency sweep inputs are used to obtain frequency responses of fixed system hub loads to a single blade IBC input. The evaluation results demonstrate the fidelity of the extracted LTI models, and thus, establish the validity of the LTI model extraction process for use in integrated flight and rotor control studies.
27
Hill, J. L. "Development of a boundary layer transition model for helicopter rotor CFD." Thesis, Cranfield University, 2005. http://dspace.lib.cranfield.ac.uk/handle/1826/10711.
Full textAbstract:
A novel transition model has been developed for use in CFD simulations of helicopter
rotor aerodynamics. The model includes significantly improved physical modelling
of the transition
processes occurring in the steady and unsteady flows found on
helicopter rotors. The model has been coupled with the k-co and k-co SST two
equation turbulence models using a novel adaptation of the technique developed by
Wilcox for the low Reynolds number k-oa model.
The method has been employed to calculate transitional flows occurring in three key
ow
regimes found in helicopter aerodynamics; that around steady and unsteady
aerofoils and that around a hovering helicopter rotor. The performance of the k-co and
the k-w SST turbulence models have been
investigated for transitional flow
simulations and the k-w SST shown to
provide substantial improvements for
transitional flows
containing separations. Dramatic improvements in the computed
pressure and skin friction distributions for several aerofoil flows have been observed
over those computed using a conventional fully turbulent simulation. Corresponding
improvements are observed in the computed lift and drag polars and transition on set is
well predicted for both low and high Reynolds number flows.
A novel structured/unstructured a priori adapted grid generation strategy has been
developed for hovering rotor flows that provides improved rotor solutions for
transitional flow analysis. The method offers vast improvements in the preservation
of vorticity in the solution at greatly reduced computational expense. Tip vortices
have been maintained to a Wake
age of 1170 degrees with just 2 million cells per
blade. The transition model has then been applied to the high quality rotor solutions
and
good agreement obtained between computed and experimental results,
highlighting that three-dimensional effects have a relatively small effect on hovering
rotor transition in-board of the blade tip. I addition, the first known verification of a
Navier-Stokes rotor code against the Fogarty semi-analytical rotating at plate case
was presented and excellent agreement obtained.
28
Lovaco, Jorge Luis. "Verication of a Modelica Helicopter Rotor Model Using Blade Element Theory." Thesis, Linköpings universitet, Fluida och mekatroniska system, 2017. http://urn.kb.se/resolve?urn=urn:nbn:se:liu:diva-142055.
Full textAbstract:
Helicopters have been valuable vehicles ever since their invention. Their capabilities for axial flight and hovering make them an outstanding resource. However, their complexity, directly related to their aerodynamics, makes them extremely hard to design. In today’s market competitivity resources must be optimized and accurate models are needed to obtain realizable designs. The well known Blade Element Theory was used to model helicopter rotors using the Modelica based software SystemModeler. However, it remained unverified due to the lack of experimental data available. The access to experimental data published by NASA motivated the comparison from the model to the measurements obtained during real testing to a scaled rotor. Some improvements were performed to the model obtaining unexpectedly accurate results for hover and axial flight. Two approaches based on the Blade Element Theory and related to Vortex Theory were followed: an infinite number of blades and a finite number of blades. Moreover, the model simulation speed was notice ably increased and prepared for the forward flight model development. Nonetheless, even though the model was ready for forward flight simulations, further research is needed due to, again, the lack of experimental data available. It is concluded from the present work that Wolfram’s SystemModeler can be used as a tool in early design phases of helicopters, even before CAD modeling and CFD due to its simplicity, speed, accuracy, and especially its capability for being used on simple desktop computers.
29
Hill, Jason Lee. "The development of a boundary layer transition model for helicopter rotor CFD." Thesis, Cranfield University, 2005. http://dspace.lib.cranfield.ac.uk/handle/1826/10711.
Full textAbstract:
A novel transition model has been developed for use in CFD simulations of helicopter rotor aerodynamics. The model includes significantly improved physical modelling of the transition processes occurring in the steady and unsteady flows found on helicopter rotors. The model has been coupled with the k-co and k-co SST two equation turbulence models using a novel adaptation of the technique developed by Wilcox for the low Reynolds number k-oa model. The method has been employed to calculate transitional flows occurring in three key ow regimes found in helicopter aerodynamics; that around steady and unsteady aerofoils and that around a hovering helicopter rotor. The performance of the k-co and the k-w SST turbulence models have been investigated for transitional flow simulations and the k-w SST shown to provide substantial improvements for transitional flows containing separations. Dramatic improvements in the computed pressure and skin friction distributions for several aerofoil flows have been observed over those computed using a conventional fully turbulent simulation. Corresponding improvements are observed in the computed lift and drag polars and transition on set is well predicted for both low and high Reynolds number flows. A novel structured/unstructured a priori adapted grid generation strategy has been developed for hovering rotor flows that provides improved rotor solutions for transitional flow analysis. The method offers vast improvements in the preservation of vorticity in the solution at greatly reduced computational expense. Tip vortices have been maintained to a Wake age of 1170 degrees with just 2 million cells per blade. The transition model has then been applied to the high quality rotor solutions and good agreement obtained between computed and experimental results, highlighting that three-dimensional effects have a relatively small effect on hovering rotor transition in-board of the blade tip. I addition, the first known verification of a Navier-Stokes rotor code against the Fogarty semi-analytical rotating at plate case was presented and excellent agreement obtained.
30
Chen, Chang. "Development of a Simplified Inflow Model for a Helicopter Rotor in Descent Flight." Diss., Georgia Institute of Technology, 2006. http://hdl.handle.net/1853/11535.
Full textAbstract:
A helicopter rotor in descent flight encounters its own wake, resulting in a doughnut-shaped ring around the rotor disk, known as the Vortex Ring State (VRS). Flight in VRS condition can be dangerous as it may cause uncommanded drop in descent rate, loss of control effectiveness, power settling, excessive thrust and torque fluctuations, and vibration. As simple momentum theory is no longer valid for a rotor in VRS, modeling of rotor inflow in VRS continues to challenge researchers, especially for flight simulation applications.
In this dissertation, a simplified inflow model, called the ring vortex model, is developed for a helicopter rotor operating in descent condition. By creating a series of vortex rings near the rotor disk, the ring vortex model addresses the strong flow interaction between the rotor wake and the surrounding airflow in descent flight. In addition, the total mass flow parameter in the existing inflow models is augmented to create a steady state transition between the helicopter and the windmill branches. With the ring vortex model, rotor inflow can now be adequately predicted over a wide range of descent rates.
Validations of the ring vortex model for helicopter rotors are conducted extensively in axial and inclined descent. Effects from blade taper, blade twist, and rotor thrust are also investigated with further application of the finite-state inflow model.
The ring vortex model is applied to a single main-rotor helicopter. The main effort is to establish VRS boundary based on heave stability criterion. In addition, two important phenomena observed in the descent flight tests are addressed in the dynamic simulation, including uncommanded drop in descent rate and loss of collective control effectiveness.
The ring vortex model is further applied to a side-by-side rotor configuration. Lateral thrust asymmetry on the side-by-side rotor configuration can be reproduced through uneven distribution of vortex rings at the two rotors. Two important issues are investigated, including the impact of vortex rings on lateral thrust deficit and on lateral AFCS limit.
31
Rodriguez, Christian. "CFD Analysis on the Main-Rotor Blade ofa Scale Helicopter Model using Overset Meshing." Thesis, KTH, Farkost och flyg, 2012. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-118797.
Full textAbstract:
In this paper, an analysis in computational uid dynamics (CFD) is presented on a helicopter scale model with focus on the main-rotor blades.The helicopter model is encapsulated in a background region and the ow eld is solved using Star CCM+. A surface and volume mesh continuum was generated that contained approximately seven million polyhedral cells, where the Finite Volume Method (FVM) was chosen as a discretization technique. Each blade was assigned to an overset region making it possible to rotate and add a cyclic pitch motion. Boundary information was exchanged between the overset and background mesh using a weighted interpolation method between cells. An implicit unsteady ow solver, with an ideal gas and a SST (Mentar) K-Omega turbulence model were used. Hover and forward cases were examined. Forward ight cases were done by changing the rotor shaft angle of attacks and the collective pitch angle 0 at the helicopter freestream Mach number of M = 0:128, without the inclusion of a cyclic pitch motion. An additional ight case with cyclic pitch motion was examined at s = 0 and = 0. Each simulation took roughly 48 hours with a total of 96 parallel cores to compute. Experimental data were taken from an existing NASA report for comparison of the results. Hover ight coincided well with the wind tunnel data. The forward ight cases (with no cyclic motion) produced lift matching the experimental data, but had diculties in producing a forward thrust. Moments in roll and pitch started to emerge. By adding a cyclic pitch successfully removed the pitch and roll moments. In conclusion this shows that applying overset meshes as a way to analyze the main-rotor blades using CFD does work. Adding a cyclic pitch motion at 0 = 5 and s = 0 successfully removed the roll and pitching moment from the results.
32
Hotchkiss, Paul. "Development of a rotor model for the numerical simulation of helicopter exterior flow-fields." Thesis, University of Cape Town, 2004. http://hdl.handle.net/11427/6774.
Full textAbstract:
Includes bibliographical references (leaves 84-85).A numerical methodology is developed to model the effect of a rotor on the surrounding flow-field. The model calculates the time-averaged aerodynamic forces exerted on the air by the fan blades within the blade-swept region, and permits the user to specify blade properties such as cross-sectional profile and orientation at a particular radial and azimuthal location. The calculated forces are included as source terms within the Reynolds-averaged Navier-Stokes equations for an incompressible fluid, which are solved by the commercial CFD solver, FLUENT. The effects of turbulence are incorporated through the use of Launder and Spalding's k-g turbulence model. This method is selected as being the most efficient use of the resources available, giving the economic advantages of a steady simulation, while allowing radial and azimuthal variations of rotor characteristics. In order to validate the accuracy of the numerical model for both aligned and non-aligned inflow conditions, results are compared with experimental data reported for an axial flow fan. Agreement between experimental and numerical results is excellent to good. Fan static pressure rise is closely predicted by the numerical solution, while fan power consumption and fan static efficiency are under and over-predicted respectively. This error may be attributed to frictional losses not accounted for in the numerical model. These include physical rotational instabilities, leading to increased mechanical losses, and tip effects due to the clearance between the fan blade tips and the fan casing. Trends are nevertheless consistently predicted by the numerical model for inflow angles up to 45°, and for the range of blade pitch settings used. The adverse effect of off-axis inflow on the fan static pressure rise is numerically predicted, while fan power consumption is found to remain independent of inflow angle, as had been experimentally observed. The rotor model is finally integrated with the fuselage of the CIRSTEL (Combined Infra-Red Suppression and Tail rotor Elimination) prototype in an analysis of the helicopter exterior flow-field. No experimental data for this configuration was available for validation purposes. However, the model is used in the simulation of several common helicopter flight conditions. Results are presented graphically, and generally indicate good agreement with physically observed phenomena.
33
Lebel, Guilhem. "Prévision des charges aéromécaniques des rotors d'hélicoptère : Application aux pales à double flèche." Thesis, Lyon, INSA, 2012. http://www.theses.fr/2012ISAL0025.
Full textAbstract:
Les récentes recherches sur les rotors d'hélicoptère conduisent au développement de pales de nouvelle génération présentant des géométries courbes. La double flèche de la pale BlueEdgeTM proposée par Eurocopter impose de reconsidérer les outils de calcul des charges rotors pour déterminer le torseur des efforts appliqués aux pales et aux éléments constitutifs du moyeu rotor afin de satisfaire aux exigences de conception et de certification. Les charges rotors se décomposent en contributions aéro- et élasto-dynamiques prises en compte par des modélisations distinctes. La thèse vise à définir une méthodologie de calcul de charges applicable aux pales à double flèche. Ainsi sont présentés les modèles aérodynamiques bi-dimensionnels pour calculer les vitesses induites du rotor et déterminer la répartition des efforts aérodynamiques sur le rotor. Le calcul des charges rotor nécessite de recourir à des modèles élasto-dynamiques. En résolvant les équations de la dynamique des solides pour un système mécanique, le code de mécanique du vol HOST considère une modélisation élastique de pale pour déterminer le torseur des efforts, les efforts de commande étant fournis par l'ensemble bielle de pas et plateaux cycliques. Le comportement non linéaire des adaptateurs de traînée interpales est décrit par des modèles de force de restitution. Ces travaux ont utilisé des caractérisations expérimentales sur des machines de traction de laboratoire ainsi que des essais en vol afin d'évaluer le niveau de représentativité des outils et méthodes proposés. La mise en oeuvre de l'ensemble de ces modèles détermine avec satisfaction les charges dynamiques du rotor pour des vols stabilisésNew generation blades have led to new load computation problems due to the evolution of the general shape, with forward and backward sweep. The BlueEdgeTM blade pattented by Eurocopter imposes to reconsider the development methodology and thus it is no longer possible to speak of straight blades and the models used for load computation have to be evaluated. The objective of this thesis is to determine what has to be modified and improved in current load computation methodology in order to reach an acceptable predictive level. This work considers both aerodynamic and dynamic models implemented in the HOST multi-body computer code. The aerodynamics models are based on the hypothesis of a two dimensional flow. The use of the CFD software \emph{elsA} is evaluated. Attention is given to rotor dynamics models that have an impact on loads, such as lead-lag damper models, blade element models and hub models. This thesis presents the different models and gives orientations relating to efficient load computation methodology. The aerodynamics models are compared to windtunnels experiments from the literature. This study leads also to perform flight tests and to investigate the dampers behavior on test benches in order to confront the computed loads to the reality of the helicopter operation. The proposed methodology is able to compute with a good accuracy rotor loads for stabilized flight cases
34
Guilbert, Bérengère. "Hybrid modular models for the dynamic study of high-speed thin -rimmed/-webbed gears." Thesis, Lyon, 2017. http://www.theses.fr/2017LYSEI127/document.
Full textAbstract:
Ces travaux de thèse ont été réalisés grâce à une collaboration entre Safran Helicopter Engines (anciennement Turbomeca) et le Laboratoire de Mécanique des Contacts et des Structures (LaMCoS) de l’INSA de Lyon (UMR CNRS 5259). Les boîtes de transmission par engrenages des moteurs d’hélicoptères convoient la puissance mécanique du turbomoteur aux accessoires (pompes, démarreur) et au rotor. Leur conception dépend des nécessités des équipements embarqués, en particulier l’allègement pour réduire la consommation en carburant. Les engrenages haute vitesse de la transmission sont allégés grâce à des enlèvements de matière dans les corps sous la denture, les voiles-minces. Un modèle dynamique d’engrenages a été développé pendant ce projet de recherche. Son approche modulaire permet l’inclusion conjointe des sollicitations dues aux vibrations de l’engrenage et de la nouvelle flexibilité des voiles-minces. Il dérive d’un modèle à paramètres concentrés, comprenant des arbres en poutre, des paliers et carters sous forme de raideurs additionnelles et un élément d’engrenage rigide inclus par son nœud central. Hypothèse est faite que tous les contacts sont situés sur les lignes de contact du plan d’action. Ces lignes sont discrétisées selon des tranches-minces dans les dents et la déviation normale des cellules est recalculée à chaque pas de temps selon la déflexion de la denture. Le nouveau modèle remplace l’engrenage rigide par une modélisation EF du pignon et/ou de la roue condensée sur les nœuds de jante. Une interface lie les raideurs du plan d’action discrétisé aux éléments finis du corps d’engrenage. L’élément prend donc en compte à la fois les sollicitations de l’engrenage et le comportement statique et modal des corps flexibles en dynamique. Des comparaisons sont faites avec des données numériques et expérimentales. Elles attestent de la capacité du nouveau modèle à prédire le comportement dynamique des engrenages flexibles à hauts régimes de rotation. Ces résultats intègrent entre autres des données locales et globales en dynamique. Finalement, le modèle est utilisé sur les deux cas académiques validés pour visualiser les effets des corps flexibles plus en détails. Un premier focus sera fait sur la déflexion statique due aux charges d’engrènement et sur l’optimisation sur le fonctionnement dynamique possible. Puis, les impacts des sollicitations de l’engrènement sur le voile en rotation seront étudiés. Enfin, le pignon et la roue seront affinés, afin de visualiser l’optimisation massique possible et son impact sur la dynamique de l’engrenageThe research work presented in this manuscript was conducted in the Contact and Structural Mechanics Laboratory (LaMCoS) at INSA Lyon, in partnership with Safran Helicopter Engines (formerly-Turbomeca). In helicopters, the power from the turboshaft is transmitted to the rotor and the various accessories (pumps, starters etc…) via transmission gearboxes. In the context of high-speed, light-weight aeronautical applications, mechanical parts such as gears have to meet somehow contradictory design requirements in terms of reliability and mass reduction thus justifying precise dynamic simulations. The present work focuses on the definition of modular gear dynamic models, capable of integrating both the local phenomena associated with the instant contact conditions between the tooth flanks and the more global aspects related to shafts, bearings and particularly the contributions of light thin-rimmed /-webbed gear bodies. The proposed models rely on combinations of condensed sub-structures, lumped parameter and beam elements to simulate a pinion-gear pair, shafts, bearings and housing. Mesh elasticity is time-varying, possibly non-linear and is accounted for by Winkler foundations derived from a classic thin-slice model. The contact lines in the base plane are therefore discretised into elemental segments which are all attributed a mesh stiffness function and a normal deviation which are updated depending on the pinion and gear angular positions. The main originality in this PhD consists in inserting condensed finite elements models to simulate flexible gear bodies while keeping the simple and faster rigid-body approach for solid gears. To this end, a specific interface has been developed to connect the discretised tooth contact lines to the continuous finite element gear body models and avoid numerical spikes in the tooth load distributions for example. A number of comparisons with numerical and experimental results show that the proposed modelling is sound and can capture most of the quasi-static and dynamic behaviour of single stage reduction units with thin-webbed gears and/or pinions. The model is then applied to the analysis of academic and industrial gears with the objective of analysing the contributions of thin, flexible bodies. Results are presented which highlight the role of centrifugal effects and tooth shape modifications at high speeds. Finally, the possibility to further improve gear web design with regard to mass reduction is investigated and commented upon
35
Tarimci, Onur. "Adaptive Controller Applications For Rotary Wing Aircraft Models Of Varying Simulation Fidelity." Master's thesis, METU, 2009. http://etd.lib.metu.edu.tr/upload/12611168/index.pdf.
Full textAbstract:
This thesis concerns the design, analysis and testing of adaptive controllers for rotary wing
aircraft, in particular helicopters. A non-linear helicopter model is developed and validated by
trim and dynamic response analyses. A inner-outer loop cascade controller is designed with
a trajectory generator in the most outer layer and an adaptive neural network controller is
implemented to the inner loop. Controller is then challenged to carry out complex maneuvers
autonomously under turbulence. Finally, the center of gravity location is varied to severe
values to observe adaptation characteristics to investigate the requirement on the knowledge
of the center of gravity location during such adaptive controller design.
36
Poyi, Gwangtim Timothy. "A novel approach to the control of quad-rotor helicopters using fuzzy-neural networks." Thesis, University of Derby, 2014. http://hdl.handle.net/10545/337911.
Full textAbstract:
Quad-rotor helicopters are agile aircraft which are lifted and propelled by four rotors. Unlike traditional helicopters, they do not require a tail-rotor to control yaw, but can use four smaller fixed-pitch rotors. However, without an intelligent control system it is very difficult for a human to successfully fly and manoeuvre such a vehicle. Thus, most of recent research has focused on small unmanned aerial vehicles, such that advanced embedded control systems could be developed to control these aircrafts. Vehicles of this nature are very useful when it comes to situations that require unmanned operations, for instance performing tasks in dangerous and/or inaccessible environments that could put human lives at risk. This research demonstrates a consistent way of developing a robust adaptive controller for quad-rotor helicopters, using fuzzy-neural networks; creating an intelligent system that is able to monitor and control the non-linear multi-variable flying states of the quad-rotor, enabling it to adapt to the changing environmental situations and learn from past missions. Firstly, an analytical dynamic model of the quad-rotor helicopter was developed and simulated using Matlab/Simulink software, where the behaviour of the quad-rotor helicopter was assessed due to voltage excitation. Secondly, a 3-D model with the same parameter values as that of the analytical dynamic model was developed using Solidworks software. Computational Fluid Dynamics (CFD) was then used to simulate and analyse the effects of the external disturbance on the control and performance of the quad-rotor helicopter. Verification and validation of the two models were carried out by comparing the simulation results with real flight experiment results. The need for more reliable and accurate simulation data led to the development of a neural network error compensation system, which was embedded in the simulation system to correct the minor discrepancies found between the simulation and experiment results. Data obtained from the simulations were then used to train a fuzzy-neural system, made up of a hierarchy of controllers to control the attitude and position of the quad-rotor helicopter. The success of the project was measured against the quad-rotor’s ability to adapt to wind speeds of different magnitudes and directions by re-arranging the speeds of the rotors to compensate for any disturbance. From the simulation results, the fuzzy-neural controller is sufficient to achieve attitude and position control of the quad-rotor helicopter in different weather conditions, paving way for future real time applications.
37
Naidoo, Vaneshen. "Implementation of a trim routine in a rotor model for the numerical simulation of helicopter flow-fields." Master's thesis, University of Cape Town, 2006. http://hdl.handle.net/11427/8911.
Full textAbstract:
Includes bibliographical references (p. 82-83)The aim of the current project is to develop, validate and implement a trim routine for a numerical rotor model, developed for the use in simulations of a helicopter exterior flow-field. In this investigation a ROBIN fuselage geometry was utilised. Simulations of the fuselage without the rotor were carried out initially so that investigations into the computational grids and turbulence models could be done. The computational simulations were performed in the commercially available CFD solver, FLUENT® Computational grids were created for the near wall modelling approach and wall function approach. Some of the more applicable turbulence models available in the solver were compared. For the wall function approach grids the k - ε, and its variants, the RNG and realizable models were found to be suitable choices. For the near wall modelling approach grids used, the SST models performed the best. The rotor model used during this investigation utilised a combination of blade element and actuator disk theory. Forces exerted by the rotor are calculated with the use of blade characteristics and flow properties. These forces were applied to the domain as momentum sources terms. The rotor model was incorporated with the CFD solver, through the use of a User Defined Function (UDF). The method used to trim the rotor was the Newton-Raphson Iterative method. This trim routine was incorporated in the UDF used for the rotor model. Tests were conducted, on a 'rotor-alone' model, as well as the rotor and fuselage model. The trim routine was found to be rigorous and managed to trim the rotor in each of the tests conducted. Good agreement between experimental and numerical collective pitch angle and cyclic pitch coefficients were found. Also the effect of the fuselage on the trim conditions proved to be minimal.
38
Samal, Mahendra Engineering & Information Technology Australian Defence Force Academy UNSW. "Neural network based identification and control of an unmanned helicopter." Awarded by:University of New South Wales - Australian Defence Force Academy. Engineering & Information Technology, 2009. http://handle.unsw.edu.au/1959.4/43917.
Full textAbstract:
This research work provides the development of an Adaptive Flight Control System (AFCS) for autonomous hover of a Rotary-wing Unmanned Aerial Vehicle (RUAV). Due to the complex, nonlinear and time-varying dynamics of the RUAV, indirect adaptive control using the Model Predictive Control (MPC) is utilised. The performance of the MPC mainly depends on the model of the RUAV used for predicting the future behaviour. Due to the complexities associated with the RUAV dynamics, a neural network based black box identification technique is used for modelling the behaviour of the RUAV. Auto-regressive neural network architecture is developed for offline and online modelling purposes. A hybrid modelling technique that exploits the advantages of both the offline and the online models is proposed. In the hybrid modelling technique, the predictions from the offline trained model are corrected by using the error predictions from the online model at every sample time. To reduce the computational time for training the neural networks, a principal component analysis based algorithm that reduces the dimension of the input training data is also proposed. This approach is shown to reduce the computational time significantly. These identification techniques are validated in numerical simulations before flight testing in the Eagle and RMAX helicopter platforms. Using the successfully validated models of the RUAVs, Neural Network based Model Predictive Controller (NN-MPC) is developed taking into account the non-linearity of the RUAVs and constraints into consideration. The parameters of the MPC are chosen to satisfy the performance requirements imposed on the flight controller. The optimisation problem is solved numerically using nonlinear optimisation techniques. The performance of the controller is extensively validated using numerical simulation models before flight testing. The effects of actuator and sensor delays and noises along with the wind gusts are taken into account during these numerical simulations. In addition, the robustness of the controller is validated numerically for possible parameter variations. The numerical simulation results are compared with a base-line PID controller. Finally, the NN-MPCs are flight tested for height control and autonomous hover. For these, SISO as well as multiple SISO controllers are used. The flight tests are conducted in varying weather conditions to validate the utility of the control technique. The NN-MPC in conjunction with the proposed hybrid modelling technique is shown to handle additional disturbances successfully. Extensive flight test results provide justification for the use of the NN-MPC technique as a reliable technique for control of non-linear complex dynamic systems such as RUAVs.
39
Panico, Pierre. "Prévision de l'amorçage de fissures de fretting par une méthode asymptotique appliquée aux rotors d'hélicoptères." Thesis, Lyon, 2019. http://www.theses.fr/2019LYSEI135.
Full textAbstract:
L'endommagement de fretting est un phénomène de fissuration apparaissant à l'interface de contacts lors de sollicitations cycliques répétées. Prévenir l’apparition de la fissuration est un enjeu majeur pour la durabilité de toutes machines tournantes et plus particulièrement les rotors d’hélicoptères. La prévision de ce phénomène dans un contexte industriel est rendu difficile par la présence de contraintes localisées. Leurs déterminations nécessitent des modèles numériques gourmands en ressources et en temps de calcul. Au cours de cette thèse, une campagne expérimentale a été réalisée sur une gamme de géométries de contact. Ces essais de fatigue en configuration de fretting ont permis de caractériser l’apparition du phénomène en fonction des conditions de contacts et de leurs sollicitations. Une approche asymptotique a été utilisée pour établir une corrélation entre les résultats expérimentaux et des propriétés matériaux connues. Cette approche consiste en une décomposition en valeurs propres du problème dans la zone critique d’amorçage de fissures de fretting. L’exploitation de ces résultats a permis la proposition d’un critère d’amorçage de fissures de fretting prenant en compte toutes les géométries de contacts ainsi que la présence de sollicitations, statiques et dynamiques. Finalement, une méthode a été développée pour appliquer ce critère d’amorçage de fissures de fretting sur des modélisations éléments finis d’assemblages industriels. Cette méthode est basée sur un enrichissement analytique autour de la singularité formée par le contact. L’application de ce critère prévoit l’apparition de l’endommagement de fretting sur tous types de géométries de manière efficace par rapports aux méthodes existantesThe fretting damage phenomenon takes place in loaded mechanical contact interfaces subject to oscillatory small displacement and could lead to cracks nucleation. The prediction of such events is a major challenge for aeronautical industry. The length scale difference between industrial models and the local stress reveals some difficulties in the risk evaluation. An experimental campaign on a wide geometrical contacting pad range has been performed. Those fatigue tests involving fretting solicitation have been used to characterize the fretting crack nucleation at 1 million cycles with various geometries and loading. A correlation between the crack nucleation results and the material properties become possible by using asymptotic parameters. They are defined as the contact edge eigenvalues and are related to the shear stress and normal stress intensity. A fretting crack nucleation criterion is proposed taking advantage of the equivalence of the asymptotic description with contact mechanics and linear elastic fracture mechanics. A numerical method is finally developed to apply this fretting criterion on finite element industrial models. The numerical tool, or analytical patch, is working as a post-processing method based on analytical stress enrichment to minimize the mesh size dependence. This analytical patch is computing the contact edge asymptotic parameters at each time step. The application of the fretting crack nucleation criterion allows a fast and robust fretting crack risk evaluation on industrial parts
40
Viswamurthy, S. R. "Piezoceramic Dynamic Hysteresis Effects On Helicopter Vibration Control Using Multiple Trailing-Edge Flaps." Thesis, 2007. https://etd.iisc.ac.in/handle/2005/561.
Full textAbstract:
Helicopters suffer from severe vibration levels compared to fixed-wing aircraft. The main source of vibration in a helicopter is the main rotor which operates in a highly unsteady aerodynamic environment. Active vibration control methods are effective in helicopter vibration suppression since they can adapt to various flight conditions and often involve low weight penalty. One such method is the actively controlled flap (ACF) approach. In the ACF approach, a trailing-edge flap (TEF) located in each rotor blade is deflected at higher harmonics of rotor frequency to reduce vibratory loads at the rotor hub. The ACF approach is attractive because of its simplicity in practical implementation, low actuation power and enhanced airworthiness, since the flap control is independent of the primary control system. Multiple-flaps are better suited to modify the aerodynamic loading over the rotor blade and hence offer more flexibility compared to a single flap. They also provide the advantage of redundancy over single-flap configuration. However, issues like the number, location and size of these individual flaps need to be addressed based on logic and a suitable performance criteria. Preliminary studies on a 4-bladed hingeless rotor using simple aerodynamic and wake models predict that multiple-flaps are capable of 70-75 percent reduction in hub vibration levels. Numerical studies confirm that multiple-flaps require significantly less control effort as compared to single-flap configuration for obtaining similar reductions in hub vibration levels. Detailed studies include more accurate aerodynamic and wake models for the rotor with TEF’s. A simple and efficient flap control algorithm is chosen from literature and modified for use in multiple-flap configuration to actuate every flap near complete authority. The flap algorithm is computationally efficient and performs creditably at both high and low forward speeds. This algorithm works reasonably well in the presence of zero-mean Gaussian noise in hub load data. It is also fairly insensitive to small changes in plant parameters, such as, blade mass and stiffness properties. The optimal locations of multiple TEF’s for maximum reduction in hub vibration are determined using Response Surface methodology. Piezoelectric stack actuators are the most promising candidates for actuation of full-scale TEF’s on helicopter rotors. A major limitation of piezoelectric actuators is their lack of accuracy due to nonlinearity and hysteresis. The hysteresis in the actuators is modeled using the classical Preisach model (CPM). Experimental data from literature is used to estimate the Preisach distribution function. The hub vibration in this case is reduced by about 81-86 percent from baseline conditions. The performance of the ACF mechanism can be further improved by using an accurate hysteresis compensation scheme. However, using a linear model for the piezoelectric actuator or an inaccurate compensation scheme can lead to deterioration in ACF performance. Finally, bench-top experiments are conducted on a commercially available piezostack actuator (APA500L from CEDRAT Technologies) to study its dynamic hysteresis characteristics. A rate-dependent dynamic hysteresis model based on CPM is used to model the actuator. The unknown coefficients in the model are identified using experiments and validated. Numerical simulations show the importance of modeling actuator hysteresis in helicopter vibration control using TEF’s. A final configuration of multiple flaps is then proposed by including the effects of actuator hysteresis and using the response surface approach to determine the optimal flap locations. It is found that dynamic hysteresis not only affects the vibration reduction levels but also the optimal location of the TEF's.
41
Viswamurthy, S. R. "Piezoceramic Dynamic Hysteresis Effects On Helicopter Vibration Control Using Multiple Trailing-Edge Flaps." Thesis, 2007. http://hdl.handle.net/2005/561.
Full textAbstract:
Helicopters suffer from severe vibration levels compared to fixed-wing aircraft. The main source of vibration in a helicopter is the main rotor which operates in a highly unsteady aerodynamic environment. Active vibration control methods are effective in helicopter vibration suppression since they can adapt to various flight conditions and often involve low weight penalty. One such method is the actively controlled flap (ACF) approach. In the ACF approach, a trailing-edge flap (TEF) located in each rotor blade is deflected at higher harmonics of rotor frequency to reduce vibratory loads at the rotor hub. The ACF approach is attractive because of its simplicity in practical implementation, low actuation power and enhanced airworthiness, since the flap control is independent of the primary control system. Multiple-flaps are better suited to modify the aerodynamic loading over the rotor blade and hence offer more flexibility compared to a single flap. They also provide the advantage of redundancy over single-flap configuration. However, issues like the number, location and size of these individual flaps need to be addressed based on logic and a suitable performance criteria. Preliminary studies on a 4-bladed hingeless rotor using simple aerodynamic and wake models predict that multiple-flaps are capable of 70-75 percent reduction in hub vibration levels. Numerical studies confirm that multiple-flaps require significantly less control effort as compared to single-flap configuration for obtaining similar reductions in hub vibration levels. Detailed studies include more accurate aerodynamic and wake models for the rotor with TEF’s. A simple and efficient flap control algorithm is chosen from literature and modified for use in multiple-flap configuration to actuate every flap near complete authority. The flap algorithm is computationally efficient and performs creditably at both high and low forward speeds. This algorithm works reasonably well in the presence of zero-mean Gaussian noise in hub load data. It is also fairly insensitive to small changes in plant parameters, such as, blade mass and stiffness properties. The optimal locations of multiple TEF’s for maximum reduction in hub vibration are determined using Response Surface methodology. Piezoelectric stack actuators are the most promising candidates for actuation of full-scale TEF’s on helicopter rotors. A major limitation of piezoelectric actuators is their lack of accuracy due to nonlinearity and hysteresis. The hysteresis in the actuators is modeled using the classical Preisach model (CPM). Experimental data from literature is used to estimate the Preisach distribution function. The hub vibration in this case is reduced by about 81-86 percent from baseline conditions. The performance of the ACF mechanism can be further improved by using an accurate hysteresis compensation scheme. However, using a linear model for the piezoelectric actuator or an inaccurate compensation scheme can lead to deterioration in ACF performance. Finally, bench-top experiments are conducted on a commercially available piezostack actuator (APA500L from CEDRAT Technologies) to study its dynamic hysteresis characteristics. A rate-dependent dynamic hysteresis model based on CPM is used to model the actuator. The unknown coefficients in the model are identified using experiments and validated. Numerical simulations show the importance of modeling actuator hysteresis in helicopter vibration control using TEF’s. A final configuration of multiple flaps is then proposed by including the effects of actuator hysteresis and using the response surface approach to determine the optimal flap locations. It is found that dynamic hysteresis not only affects the vibration reduction levels but also the optimal location of the TEF's.
42
Murugan, M. Senthil. "Aeroelastic Analysis And Optimization Of Composite Helicopter Rotor With Uncertain Material Properties." Thesis, 2009. https://etd.iisc.ac.in/handle/2005/1117.
Full textAbstract:
Incorporating uncertainties in the aeroelastic analysis increases the confidence levels of computational predictions and reduces the need for validation with experimental or flight test data. Helicopter rotor blades, which play a dominant role in the overall vehicle performance, are routinely made of composites. The material properties of composites are uncertain because of the variations in manufacturing process and other effects while in service, maintenance and storage. Though nominal values are listed, they are seldom accurate. In this thesis, the effect of uncertainty in composite material properties on the computational predictions of cross-sectional properties, natural frequencies, blade tip deflections, vibratory loads and aeroelastic stability of a four-bladed composite helicopter rotor is studied.
The effect of material uncertainty is studied with the composite rotor blades modeled as components of soft-inplane as well as stiff-inplane hingeless helicopter rotors. Aeroelastic analysis based on finite elements in space and time is used to evaluate the helicopter rotor blade response in hover and forward flight. Uncertainty analysis is performed with direct Monte Carlo simulations based on a sufficient number of random samples of material properties. It is found that the cross-sectional stiffness parameters and natural frequencies of rotor blades show considerable scatter from their baseline predictions. The uncertainty impact on the rotating natural frequencies depends on the level of centrifugal stiffening of each mode. The propagation of material uncertainty into aeroelastic response causes large deviations from the baseline predictions. The magnitudes of 4/rev vibratory loads show deviations of 10 to 600 percent from their baseline predictions. The aeroelastic stability in hover and forward flight conditions also show considerable uncertainty in the predictions. In addition to the effects of material uncertainty, various factors influencing the propagation of material uncertainty are studied with the first-order based reliability methods. The numerical results have shown the need to consider the uncertainties in the helicopter aeroelastic analysis for reliable computational predictions.
Uncertainty quantification using direct Monte Carlo simulation is accurate but computationally expensive. The application of response surface methodologies to reduce the computational cost of uncertainty analysis is studied. Response surface approximations of aeroelastic outputs are developed in terms of the composite material properties. Monte Carlo simulations are then performed using these computationally less expensive response surface models. The results of this study show that the metamodeling techniques can effectively reduce the computational cost of uncertainty analysis of composite rotor blades.
In the last part of the thesis, an aeroelastic optimization method to minimize the vibration level is developed with due consideration to material uncertainty. Second-order polynomial response surfaces are used to approximate the objective function which smooths out the local minima or numerical noise in the design space. The aeroelastic optimization is carried out with the nominal values of composite material properties and the performance of final design is found to be optimum even for the perturbed values of material properties.
43
Murugan, M. Senthil. "Aeroelastic Analysis And Optimization Of Composite Helicopter Rotor With Uncertain Material Properties." Thesis, 2009. http://hdl.handle.net/2005/1117.
Full textAbstract:
Incorporating uncertainties in the aeroelastic analysis increases the confidence levels of computational predictions and reduces the need for validation with experimental or flight test data. Helicopter rotor blades, which play a dominant role in the overall vehicle performance, are routinely made of composites. The material properties of composites are uncertain because of the variations in manufacturing process and other effects while in service, maintenance and storage. Though nominal values are listed, they are seldom accurate. In this thesis, the effect of uncertainty in composite material properties on the computational predictions of cross-sectional properties, natural frequencies, blade tip deflections, vibratory loads and aeroelastic stability of a four-bladed composite helicopter rotor is studied.
The effect of material uncertainty is studied with the composite rotor blades modeled as components of soft-inplane as well as stiff-inplane hingeless helicopter rotors. Aeroelastic analysis based on finite elements in space and time is used to evaluate the helicopter rotor blade response in hover and forward flight. Uncertainty analysis is performed with direct Monte Carlo simulations based on a sufficient number of random samples of material properties. It is found that the cross-sectional stiffness parameters and natural frequencies of rotor blades show considerable scatter from their baseline predictions. The uncertainty impact on the rotating natural frequencies depends on the level of centrifugal stiffening of each mode. The propagation of material uncertainty into aeroelastic response causes large deviations from the baseline predictions. The magnitudes of 4/rev vibratory loads show deviations of 10 to 600 percent from their baseline predictions. The aeroelastic stability in hover and forward flight conditions also show considerable uncertainty in the predictions. In addition to the effects of material uncertainty, various factors influencing the propagation of material uncertainty are studied with the first-order based reliability methods. The numerical results have shown the need to consider the uncertainties in the helicopter aeroelastic analysis for reliable computational predictions.
Uncertainty quantification using direct Monte Carlo simulation is accurate but computationally expensive. The application of response surface methodologies to reduce the computational cost of uncertainty analysis is studied. Response surface approximations of aeroelastic outputs are developed in terms of the composite material properties. Monte Carlo simulations are then performed using these computationally less expensive response surface models. The results of this study show that the metamodeling techniques can effectively reduce the computational cost of uncertainty analysis of composite rotor blades.
In the last part of the thesis, an aeroelastic optimization method to minimize the vibration level is developed with due consideration to material uncertainty. Second-order polynomial response surfaces are used to approximate the objective function which smooths out the local minima or numerical noise in the design space. The aeroelastic optimization is carried out with the nominal values of composite material properties and the performance of final design is found to be optimum even for the perturbed values of material properties.
44
Cohen, Gary M. "A design study of a scale model bearingless helicopter rotor system using composite materials." Thesis, 1991. http://hdl.handle.net/10539/22844.
Full textAbstract:
A thesis submitted to the Faculty of Engineering, University of the
Witwatersrand, Johannesburg, in fulfillment of the requirements for the
degree of Master of Science in Engineering.The use of advanced composite materials in helicopter rotor systems offers opportunities for improvements in aerodynamic geometry, performance, weight, damage tolerarice, maintenance and operating costs. Technical aspects of the design and analysis and the-practical aspects of the manufacture of a composite rotor system are discussed herein. The rotor system was compared to an existing conventional teetering rotor system, in order to establish the viability of the new composite rotor system, This rotor system reduced the number of components by 55% and the manufacturing time by half, due to the simplicity of the design and lay up procedure, thus making the system economically more viable. The mass was predicted to within 1% of that achieved in practice and gave a mass advantage of 50.5% over the conventional rotor. Static tests identified the failure modes and stress concentration points, while. the comparative hover tests showed the system to have ±20% less drag. [Abbreviated Abstract. Open document to view full version]
AC2017
45
Lin, Jang-Hau, and 林章豪. "Rotor Speed Control for Remote Control Model Helicopter." Thesis, 2004. http://ndltd.ncl.edu.tw/handle/06263789088809918258.
Full textAbstract:
碩士元智大學
機械工程學系
92
We design a rotor speed controller for R/C model helicopter to limit the vibration and jitter due to the unstable lift and thrust force caused by main rotor under varying rotor speed. We implement the CAN bus to regulate the rotor speed as CAN bus has real time protocol and multi processor function. Two Microchip PIC18F458 are installed in our model helicopter. One Microchip PIC18F458 measures the engine speed from the tail rotor, and the other control the servo motor to actuate the fuel valve. DS1103 of dSPACE is a signal processor to collect the data and control the helicopter in manual mode through the CAN bus protocol. In our control system, we focus the fuel valve actuation and tail rotor speed to design a system identification. RST controller with anti-windup function has been test successfully with real loading and disturbance imposed, and the differential rotor speed reduced to 2.5 rpm compare to 7.3 rpm under open loop control.
46
Chia-Sheng, Liu, and 劉家昇. "Attitude Control of a Two Rotor Helicopter Model with LQR Control." Thesis, 2005. http://ndltd.ncl.edu.tw/handle/35006009695436149265.
Full textAbstract:
碩士國立臺灣科技大學
電機工程系
93
In this thesis, the mathematical model and LQR control of a two rotors helicopter model system is studied. The mathematical model of the twin rotor MIMO helicopter system is derived from Lagrange dynamics, and it’s propeller thrusts are analyzed. Based on the linearized model, a PID controller and LQR controller are designed for stabilization control, disturbance rejection and tracking control. Computer simulations and experimental results are performed to check the feasibility of the proposed linear control methods.