Dissertations / Theses on the topic 'Airfoil profiles'

An experimental study has been carried out on an airfoil with slotted flap and spoiler using Joukowsky profiles. Pressure distributions were measured as functions of angle of attack, flap deflection angle, spoiler size and inclination. The results are uncorrected for wind tunnel wall effects but the data base is available to carry out the corrections. The results will be used to compare with predictions of a theoretical model, yet to be worked out, which combines work previously done by Williams, Jandali, Parkinson and Yeung. This theory will involve the potential flow about a two-element 'near'-Joukowsky airfoil system. The secondary airfoil is a simulated slotted flap. Various size spoilers are introduced to the system at arbitrary angles of inclination using methods proposed by Parkinson and Yeung. The experimental results are qualitatively reasonable and some interesting effects are observed. The behaviour of spoilers, when used with slotted flaps at various deflection angles, corresponds well with requirements of aircraft in approach or landing situations. Similarly, the use of slotted flaps alone provides the high lift at low angle of attack which is beneficial to aircraft taking off. Some recommendations are proposed for further testing with this equipment.
Applied Science, Faculty of
Mechanical Engineering, Department of
Graduate

The growing demand for ever more greener aero engines has led to ever more challenging designs and higher quality products. An investigation into Coordinate Measuring Machine measurement uncertainty using physical measurements and virtual simulations revealed that there were several factors that can affect the measurement uncertainty of a specific task. Measurement uncertainty can be affected by temperature, form error and measurement strategy as well as Coordinate Measuring Machine specification. Furthermore the sensitivity of circular features size and position varied, when applying different substitute geometry algorithms was demonstrated. The Least Squares Circle algorithm was found to be more stable when compared with the Maximum Inscribed Circle and the Minimum Circumscribed Circle. In all experiments it was found that the standard deviation when applying Least Squares Circle was of smaller magnitude but similar trends when compared with Maximum Inscribed Circle and the Minimum Circumscribed Circle. A Virtual Coordinate Measuring Machine was evaluated by simulating physical measurement scenarios of different artefacts and different features. The results revealed good correlation between physical measurements uncertainty results and the virtual simulations. A novel methodology for the automated assessment of leading edge airfoil profiles was developed by extracting the curvature of airfoil leading edge, and the method lead to a patent where undesirable features such as flats or rapid changes in curvature could be identified and sentenced. A software package named Blade Inspect was developed in conjunction with Aachen (Fraunhoufer) University for the automated assessment and integrated with a shop floor execution system in a pre-production facility. The software used a curvature tolerancing method to sentence the leading edge profiles which aimed at removing the subjectivity associated with the manual vision inspection method. Initial trials in the pre-production facility showed that the software could sentence 200 profiles in 5 minutes successfully. This resulted in a significant improvement over the current manual visual inspection method which required 3 hours to assess the same number of leading edge profiles.

This report concerns the research, design, implementation and application of an airfoil trailing-edge noise prediction tool in the development of new, quieter airfoil for large-size wind turbine application. The tool is aimed at enabling comparative acoustic performance assessment of airfoils during the early development cycle of new blades and rotors for wind turbine applications. The ultimate goal is to enable the development of quieter wind turbines by the Wind Energy Industry. The task was accomplished by developing software that is simultaneously suitable for comparative design, computationally efficient and user-friendly. The tool was integrated into a state-of-the-art wind turbine design and analysis code that may be downloaded from the web, in compiled or source code form, under general public licensing, at no charge. During the development, an extensive review of the existing airfoil trailing-edge noise prediction models was accomplished, and the semi-empirical BPM model was selected and modified to cope with generic airfoil geometry. The intrinsic accuracy of the original noise prediction model was evaluated as well as its sensitivity to the turbulence length scale parameter, with restrictions imposed accordingly. The criterion allowed comparison of performance of both CFD-RANS and a hybrid solver (XFLR5) on the calculation of the turbulent boundary layer data, with the eventual adjustment and selection of the latter. After all the elements for assembling the method had been selected and the code specified, a collaboration project was made effective between Poli-USP and TU-Berlin, which allowed the seamless coupling of the new airfoil TE noise module, \"PNoise\", to the popular wind turbine design/analysis integrated environment, \"QBlade\". After implementation, the code calculation routines were thoroughly verified and then used in the development of a family of \"silent profiles\" with good relative acoustic and aerodynamic performance. The sample airfoil development study closed the initial design cycle of the new tool and illustrated its ability to fulfill the originally intended purpose of enabling the design of new, quieter blades and rotors for the advancement of the Wind Energy Industry with limited environmental footprint.
Este trabalho descreve a pesquisa de elementos iniciais, o projeto, a implantação e a aplicação de uma ferramenta de predição de ruído de bordo de fuga, no desenvolvimento de aerofólios mais silenciosos para turbinas eólicas de grande porte. O objetivo imediato da ferramenta é permitir a comparação de desempenho acústico relativo entre aerofólios no início do ciclo de projeto de novas pás e rotores de turbinas eólicas. O objetivo mais amplo é possibilitar o projeto de turbinas eólicas mais silenciosas, mas de desempenho aerodinâmico preservado, pela indústria da Energia Eólica. A consecução desses objetivos demandou o desenvolvimento de uma ferramenta que reunisse, simultaneamente, resolução comparativa, eficiência computacional e interface amigável, devido à natureza iterativa do projeto preliminar de um novo rotor. A ferramenta foi integrada a um ambiente avançado de projeto e análise de turbinas eólicas, de código aberto, que pode ser livremente baixado na Web. Durante a pesquisa foi realizada uma ampla revisão dos modelos existentes para predição de ruído de bordo de fuga, com a seleção do modelo semi-empírico BPM, que foi modificado para lidar com geometrias genéricas. A precisão intrínseca do modelo original foi avaliada, assim como sua sensibilidade ao parâmetro de escala de turbulência transversal, com restrições sendo impostas a esse parâmetro em decorrência da análise. Esse critério permitiu a comparação de resultados de cálculo provenientes de método CFD-RANS e de método híbrido (XFLR5) de solução da camada limite turbulenta, com a escolha do último. Após a seleção de todos os elementos do método e especificação do código, uma parceria foi estabelecida entre a Poli-USP e a TU-Berlin, que permitiu a adição de um novo módulo de ruído de bordo de fuga, denominado \"PNoise\", ao ambiente de projeto e análise integrado de turbinas eólicas \"QBlade\". Após a adição, as rotinas de cálculo foram criteriosamente verificadas e, em seguida, aplicadas ao desenvolvimento de aerofólios mais silenciosos, com bons resultados acústicos e aerodinâmicos relativos a uma geometria de referência. Esse desenvolvimento ilustrou a capacidade da ferramenta de cumprir a missão para a qual foi inicialmente projetada, qual seja, permitir à Indústria desenvolver pás mais silenciosas que irão colaborar com o avanço da energia eólica através da limitação do seu impacto ambiental.

The background for this thesis originates from a study of the flow characteristics for an airfoil of the type NACA0018. The aim for this thesis was to evaluate how the characteristics of the flow over the NACA0018 profile depend on surface roughness. Airfoils were manufactured in Aluminum by Computer Numerical Control-milling and in polylactide polymer using a 3D-printer, where some of the profile surfaces were postprocessed with sandpaper in various grain sizes. The surface roughness of the profiles was evaluated in a 3D optical profilometer using white light interferometry from Filmetrics. By that technique 3D surface plots were created. The manufactured airfoils were tested in a wind tunnel where the achieved data was made dimensionless for comparative purposes. The computational fluid dynamics simulations were performed in Ansys Fluent and compared against the wind tunnel data as well as with the data from a previously made study at htw saar. The results from the wind tunnel tests show that the surface roughness has an effect on the flow characteristic of the airfoil, where different angles of stall were observed in the comparison. The difference for the dimensionless numbers coefficient of lift and drag show that the manufactured aluminum airfoil performs better compared to the 3D-Printed airfoil in this study. It has a higher performance mean value for both of these coefficients in a span of angles between 0 and 30 degrees. When compared, the results from the simulations and wind tunnel experiments do match in some cases, where the dimensionless coefficients and stall angle coheres. Further studies based on this report are recommended, where small geometric changes to the profile could be tested and validated.

This thesis is focused on the aerodynamic analysis of the competition paraglider wing. Drags of the particular wing parts are divided into chapters. The aim was to get a grasp of sizes of the individual components drags in relation to the entire assembly. In the first instance, a 2D profile and then the entire configuration of the 3D wing was analyzed. After the evaluation, some power reserves were detected in an airfoil and so the airfoil shape was optimized. After the optimization of the individual components, the CFD calculation was used again. At the end, geometry changes were evaluated.

Kenntnisse über Parametereigenschaften und Charakteristiken von Flugzeugkomponenten sind eine wesentliche Grundlage für Methoden des Flugzeugentwurfs. Daher ist Ziel dieser Arbeit, statistische Merkmale und Kenngrößen einer für den Flugzeugbau und Entwurf relevanten Auswahl an Komponenten zu erschließen. Dabei wurden zunächst superkritische Tragflügelprofile hinsichtlich ihrer geometrischen Eigenschaften (relative Profildicke, Wölbung, Dickenrücklage, Wölbungsrücklage und der sogenannte "Leading Edge Sharpness Parameter") untersucht. Diese Eigenschaften wurden mit der Software XFLR5 aus einer Auswahl an superkritischen Profilgeometrien erhoben und mit grafischen und beschreibenden Statistikmethoden ausgewertet. Die Profile wiesen relative Wölbungen von 0 % bis 3,4 % auf, die Mehrzahl entfiel auf Wölbungen von 1 % bis 2 %. Die Wölbungsrücklagen zeigten die für superkritische Profile typische Lage im hinteren Profilbereich zwischen 70 % und 90 % der Profiltiefe. Die Dickenrücklagen verteilten sich um einen Mittelwert von 37 % der Profiltiefe. Eine Betrachtung von Flugzeugreifendimensionen sollte das Verhältnis von Reifenbreite zum Durchmesser w/d charakterisieren. Es wurde ein annähernd lineares Verhalten festgestellt. Die Werte des Parameters w/d umfassten einen Bereich von 0,3 bis 0,4. Durch Regressionsanalysen konnten auch die Abhängigkeiten des Parameters w/d von nur einer bekannten Reifendimension (Breite oder Durchmesser) aufgezeigt werden. Die im Rahmen dieser Arbeit dargestellten Erkenntnisse können als Grundlage weiterführender Untersuchungen genutzt werden.

Le bruit tonal rayonné au bord de fuite des profils à faible nombre de Reynolds est un phénomène observé sur les ailes de drones ou micro-drones qui sont utilisés partout dans la vie quotidienne. La diminution de ce bruit va augmenter la survivabilité et l'efficacité des appareils dans le domaine militaire. De plus, cela va augmenter le champ des applications civiles et minimiser la pollution par le bruit. La réduction efficace du bruit est indispensable et, par conséquent, une compréhension complète du processus de rayonnement du bruit tonal du profil est nécessaire. Malgré le fait que des essais dédiés aient été réalisés depuis les années 70, il reste beaucoup de détails à expliquer. Le travail présenté est dédié à une étude expérimentale et analytique du bruit tonal. C'est une partie de collaboration entre l'Ecole Centrale de Lyon et Embry- Riddle Aeronautical University. Le but est de réaliser une caractérisation exhaustive des paramètres acoustiques et aérodynamiques du bruit tonal de bord de fuite d'un profil et de produire une base de données qui pourra être utilisée pour valider les simulations numériques réalisées dans le futur. Le profil symétrique NACA-0012 ainsi que le profil asymétrique SD7003 ont été testés pour une série d'angles d'incidence (de -10° à 10°) dans la soufflerie anéchoïque à jet ouvert de l'Ecole Centrale de Lyon pour des nombres de Reynolds modérés (0.6x105 < Rec < 2.6x105). Les mesures de pression aux parois et de pression acoustique en champ lointain pour différentes configurations ont permis d'observer une structure en escalier de la signature du bruit, de déterminer quelle face du profil a produit le bruit et de distinguer le rôle de la boucle de rétroaction. Des techniques supplémentaires de post-traitement comme l'analyse temps-fréquence ont montré l'existence de plusieurs régimes (un régime de commutation entre deux états, un régime d'une seul fréquence et un régime à plusieurs fréquences) de l'émission de bruit. L'analyse de bi-cohérence a montré qu'il y a des couplages nonlinéaires entre les fréquences. Une étude par l'anémométrie à fil chaud et par des techniques de visualisation de l'écoulement a montré que la formation d'une bulle de décollement est une condition nécessaire mais pas suffisante pour la génération du bruit. De plus, la localisation de la bulle est aussi importante et elle doit être suffisamment proche du bord de fuite. En outre, l'analyse de stabilité linéaire des résultats de simulations numériques a montré que des ondes de Tollmien-Schlichting sont transformées en ondes de Kelvin-Helmholtz dans la zone du décollement. Une prédiction analytique de l'amplitude des fréquences pures émises dans le champ lointain a été effectuée sur la base du modèle d'Amiet en supposant que le champ de pression pariétal est bidimensionnel. Les mesures de pression proches du bord de fuite du profil ont été prises comme données d'entrée. Les amplitudes prédites sont globalement en accord avec les mesures acoustiques. Après l'analyse de tous les résultats la description suivante du processus de rayonnement de sons purs peut être proposée. Les ondes de Tollmien-Schlichting qui se développent initialement dans la couche limite se transforment en ondes de Kelvin-Helmholtz le long de la couche de cisaillement de la bulle de décollement. Au bord de fuite du profil elles sont converties en ondes acoustiques qui forment un couplage fort avec les instabilités de couche limite plus en amont de l'écoulement, pilotant elles-mêmes le déclenchement de ces instabilités
The tonal trailing-edge noise generated by transitional airfoils is a topic of interest because of its wide area of applications. One of them is the Unmanned Air Vehicles operated at low Reynolds numbers which are widely used in our everyday life and have a lot of perspectives in future. The tonal noise reduction will increase the survivability and effectiveness of the devices in military field. Moreover it will enlarge the range of civil use and minimize noise pollution. The effective noise reduction is needed and therefore the complete understanding of the tonal noise generation process is necessary. Despite the fact that investigation of the trailing-edge noise was started since the seventies there are still a lot of details which should be explained. The present work is dedicated to the experimental and analytical investigation of the tonal noise and is a part of the collaboration project between Ecole Centrale de Lyon and Embry-Riddle Aerospace University. The aim is to conduct an exhaustive experimental characterization of the acoustic and aerodynamic parameters of the trailing-edge noise and to produce a data base which can be used for further numerical simulations conducted at Embry-Riddle Aerospace University. A symmetric NACA-0012 airfoil and a slightly cambered SD7003 airfoil at moderate angles of attack (varied from -10° à 10°) were tested in an open-jet anechoic wind tunnel of Ecole Centrale de Lyon at moderate Reynolds numbers (0.6x105 < Rec < 2.6x105). Measurements of the wall pressure and far-field acoustic pressure in different configurations allowed to observe the ladder-type structure of the noise signature, to determine which side produced tones and to distinguish the role of the acoustic feedback loop. Additional post-processing techniques such as time-frequency analysis showed the existence of several regimes (switching regime between two tones, one-tone regime and multiple-tones regime) of noise emission. The bicoherence analysis showed that there are non-linear relationships between tones. The investigation of the role of the separation area by hot-wire anemometry and flow visualization techniques showed that the separation bubble is a necessary but not a suficient condition for the noise generation. Moreover the location of the bubble is also important and should be close enough to the trailing edge. Furthermore the linear stability analysis of accompanying numerical simulation results showed that the Tollmien-Schlichting waves transform to the Kelvin-Helmholtz waves at the separation area. An analytical prediction of the tone levels in the far-field was done using Amiet's model based on the assumption of perfectly correlated sources along the span. The wall-pressure measurements close to the trailing edge were used as an input data. The comparisons of the predicted levels and measured ones showed a good agreement. After analysis of all results the following description of the tonal noise mechanism is proposed. At some initial point of the airfoil the Tollmien-Schlichting instabilities start. They are traveling downstream and continued to Kelvin-Helmholtz waves along the shear-layer of the separation bubble. These waves reach the trailing edge, scatter from it as acoustic waves, which move upstream. The acoustic waves amplify the boundary layer instabilities at some frequencies for which the phases of both motions match and creates the feedback loop needed to sustain the process

The thesis contains determination of real aerodynamics characteristics of PW09-135 flap airfoil with help of calibration method via CFD, followed by accomplishment of CFD analysis of parts and the whole Twin Shark glider. Detailed research and optimization was made to flow field quality of fuselage and wing fuselage junction. On the basis of CFD solution was determine a stability control, static margin and angle of attack of a horizontal stabilizer. The thesis concludes an evaluation of results applicability in praxis.

The main objective of this diploma thesis is to optimize the high lift device on the wing of the Phoenix Air U-15 ultralight aircraft, so that it complies with the UL-2 regulation regarding the stalling speed – 65 KPH. This is fulfilled by optimization of the slotted flap position. Methods used include the Response Surface Method and the Computational Fluid Dynamics approach – namely Ansys Fluent v6 software package. Furthermore, the paper deals with take-off flap optimization and construction of the flap deflection mechanism.

The work deals with the use of geometric parameterization for shape description of some parts of the airplane. Geometric parametrization is used for creating a parametric model airfoil. This parametric model allows local deformations pobrchu profile and can easily be applied to generate the geometry of the wing or other parts letoumu. Some properties of the parametric model were tested applications in aerodynamic optimization. Furthermore, the work deals with the parametric description of the blades, the aerodynamic optimization and noise analysis. For propeller blade were created distribution function of the control parameters that can be used in aerodynamic optimization of the blades. Geometric parameterization is used for identifying the location and other characteristics of noise sources.

Ce mémoire de thèse est consacré à l'étude d'un schéma numérique permettant, à terme, d'effectuer des calculs Navier-Stokes précis à l'aide d'une intégration Euler implicite non-factorisée. La discrétisation en espace est celle d'un schéma décentré hybride (HUS), d'après les travaux de Coquel & Liou, qui conjugue le schéma de Perthame (FVS), fondé sur une interprétation cinétique des équations d'Euler, et le schéma de Osher (FDS). Le schéma ainsi bati préserve la robustesse des FVS, au travers des ondes non-linéaires, et la précision des FDS, pour la capture des ondes linéaires. L'opérateur implicite est construit par linéarisation du FVS à l'ordre un en espace et par linéarisation des conditions-limites aussi bien aux frontières physiques qu'aux frontières de raccord multi-domaines. Le système linéaire qui en résulte est inversé de manière approchée à l'aide d'un algorithme itératif de Krylov, ici Bi-CGStab dû à Van der Vorst, que l'on préconditionne par un autre algorithme itératif SSOR. La validation de l'opérateur spatial est effectuée sur des cas de tube à choc et sur un profil d'aile par comparaison avec le schéma de Jameson. Ensuite, l'intégration temporelle implicite est évaluée à l'aide de cas bidimensionnels, un canal avec bosse et le même profil d'aile, et à l'aide de cas tridimensionnels, une nacelle isolée et une configuration comprenant un tronçon d'aile, un mât et une nacelle perméable. Ces études mettent en évidence, d'une part, la nette amélioration de la représentation des sillages et la réduction des pertes de pression totale obtenue à l'aide du schéma hybride par rapport à un FVS seul et, d'autre part, le bien fondé de l'implicitation des conditions-limites, de l'augmentation forcée du CFL, d'une résolution seulement approchée du système linéaire et du préconditionnement itératif
This thesis dissertation aims at studying a numerical scheme that allows accurate Navier-Stokes computations using an Euler implicit non-factorised integration. The spatial operator has been constructed by the combination of the scheme of Perthame (FVS), based on a kinetic interpretation of the Euler equations, and of the scheme of Osher (FDS). This results in one member of the hybrid upwind scheme family (HUS) introduced by Coquel & Liou, which retains the FVS robustness through non-linear waves and the FDS accuracy in capturing linear waves. The implicit operator is provided by the exact linearization of the FVS part, at first order in space, and by incorporating the boundary conditions, inter-block matching ones as well as physical ones. The subsequent linear system is approximately inverted by means of a Krylov algorithm, Bi-CGStab from Van der Vorst, and is preconditioned by an SSOR iterative method. The spatial discretization has been validated on shock tube problems and on an aerofoil profile by comparison with the scheme of Jameson. Then, the implicit integration has been assessed on bi-dimensional cases, a channel with a bump and the same aerofoil profile, and on three-dimensional cases, an isolated nacelle and a full configuration made of a truncated wing, a pylon and a nacelle. Numerical tests have demonstrated, on the one hand, a serious improvement in capturing wakes and in reducing total pressure losses, due to the hybrid scheme over a simple FVS scheme and, on the other hand, the relevance of implicit boundary conditions, the constant increase of the CFL number, a mere approximate inversion of the linear system and the efficiency of the iterative preconditioner

博士
國立中央大學
機械工程研究所
95
The blade is the most important component in aircraft engines while its geometry and the corresponding manufacturing quality significantly affect the engine’s performance and life time. Therefore, the inspection is very important procedure for quality assurance. However, the reconstruction of the datum planes and the datum points referring to the blueprint and the manufacturing machine is difficult for the initialization of the inspection procedure. The traditional methods are based the operator’s experience and relied on the manually operations. This is not only time consuming but also lacking accuracy. The coordinate reconstruction is very difficult to achieve in this way. This research categorizes several procedures from most kind blades from blueprint to determine the datum points and datum planes for automatic inspection. The purpose of these proposed procedures is to reconstruct the inspection coordinate system as that in the manufacturing stage. Both of three-plane method and six-point iteration method may apply to rebuild the manufacturing coordinate system in the inspection procedure. The blade inspection focuses on the airfoil profile and the error analysis in the key dimensions. However, the traditional measurement methods may not sufficient to handle the inspection of the sequence of airfoil profile due to the blade design shown in most of the blade blueprint. This study proposed the off-line inspection procedure regarding to the airfoil profile measurement while the original CAD model was provided. A two-stage inspection procedure is also proposed in order to examine the blade with the blueprint only. Regarding to the significant change of the curvature rate in the leading or trailing edges of the airfoil, a searching and measurement method is also proposed to reduce the non-smooth measurements. It is very difficult to calculate the airfoil parameters and analyze the profile error by the specification in blueprint and the profile measurement data since the uncertainty in the datum planes and datum points. Therefore, based on the objective of automatic inspection and then calculate the airfoil parameters and the profile error analysis. The proposed optimum datum determination method can be achieved by iteration in searching the shortest measurement distance and calculate the transform matrix for coordinate system correction. Regarding to blades inspection, this study proposed a complete automatic procedure includes (1) Blade datum determination methods for several kinds of blades, (2) Rebuild the manufacturing coordinate system methods, (3) Airfoil profile measurement methods, (4) Airfoil analysis methods, and (5) Automatic inspection methods. The proposed procedure and the corresponding inspection techniques were proved the capabilities in several aero engine blades and successfully applied by several domestic aero industry companies for the blade inspection and analysis.

Title: Solution of inverse problem for a flow around an airfoil Author: Mgr. Jan Šimák Department: Department of Numerical Mathematics Supervisor: prof. RNDr. Miloslav Feistauer, DrSc., dr. h. c., Department of Numerical Mathematics Abstract: The method described in this thesis deals with a solution of an inverse problem for a flow around an airfoil. It can be used to design an airfoil shape according to a specified velocity or pressure distribution along the chord line. The method is based on searching for a fixed point of an operator, which combines an approximate inverse and direct operator. The approximate inverse operator, derived on the basis of the thin airfoil theory, assigns a corresponding shape to the specified distribution. The resulting shape is then constructed using the mean camber line and thickness function. The direct operator determines the pressure or velocity distribution on the airfoil surface. We can apply a fast, simplified model of potential flow solved using the Fredholm integral equation, or a slower but more accurate model of RANS equations with a k-omega turbulence model. The method is intended for a subsonic flow.

Rotating blades, which are the most critical components of any turbo-machinery, need to be designed to withstand forced vibrations due to accidental tip rub impact against inner surface of casing. These vibrations are typically dependent on operating conditions and geometric parameters. In the current study, a rotor test rig with a maximum tip speed capability of 144 km/hr has been developed for studying the dynamic behavior of representative jet engine compressor blades actuated by the closure of clearance between the tip of a given rotating blade and a sector of the inner lining of the casing. Ten different blade profiles are chosen in the present research. The blades are obtained by lofting NACA GOE123 airfoil cross-section along different stacking axes. Rotor test rigs which simulate transient dynamic events require high frequency data acquisition systems like slip ring arrangement or telemetric transmission. While slip rings introduce noise into the signal, the telemetric transmission works out to be rather expensive. To circumvent the stated shortcomings of data acquisition systems, a novel rotor-mounted data acquisition system has been implemented here which captures dynamic strains in vibrating blades during operation. The current data acquisition system can store data for duration of five seconds with a sampling rate of 35 kHz. It has been calibrated with four standard tests, and provides a simple and efficient mode of data capturing. Three blades with airfoil sections (a flat beam-type blade of uniform rectangular cross-section, a blade with twisted cross-sections stacked along a straight line, and a blade similar to the latter but with a curved stacking axis) are tested under controlled rub conditions at four different speeds. The maximum test speed is restricted to 800 rpm for reasons of safety although the set-up is designed to operate up to a maximum speed of 2000 rpm. For each of the rotor speeds, a blade is tested for three to four different stagger angles in the range of 0o-30o. By plotting the RMS values of measured dynamic responses with respect to stagger angle for a given rotor speed, it has been observed, perhaps for the first time in published literature, that a stagger angle of around 20o yields the maximum RMS value of strain response. A major objective of the current study has been to utilize the data generated in the tip rub impact tests for validating a predictive numerical model of the test set-up using explicit finite element analysis. To this end, a finite element model of the rotor rig inclusive of a rotor with two blades and the static frame structure is developed and analyzed using an explicit LS-DYNA solver. This model is calibrated with the test results of the three blade designs described above. In particular, it has been shown that the frequency contents of the measured dynamic strain responses agree quite well with frequencies obtained from the numerically computed responses. It has been found in the experimental responses that a given blade vibrates with two main frequencies: one corresponding to the first natural frequency of the rotor-blade system during the tip-rubbing phase (which lasts until the blade tip is in contact with the rub element which is a sector of the circular casing), and another corresponding to the first natural frequency of the blade when it vibrates freely without its tip being in contact with the rub-liner of the casing. A shortcoming of the current modeling approach is its inability to realistically represent the damping behaviors observed in the tests. For reasons of computational efficiency and consistent with the fact that there was no perceptible damage in the tested blades, an elastic constitutive behavior is specified for the blades, while the sacrificial PVC rub-liner is assumed to behave elasto-plastically. A limited study has also been carried out by assigning an elasto-plastic constitutive model to one of the blades previously represented with elastic properties only, and although incipient yielding is observed in a highly localized region at the tip of a blade (which can also be a numerical artifact), the responses under the two material behavior considerations (i.e. elastic and elasto-plastic) are found to be nearly same. Finally, this validated modeling approach is applied to the study of blades of ten distinct geometric profiles (including the three configurations already considered) at a speed of 800 rpm and the resonant speed of a given blade. Comparisons are made between the relevant responses (such as time-histories of root strain, shaft torque, blade axial displacement, bearing load and rub force) of nine blades with airfoil cross-sections (leaving aside the results for the first blade of rectangular cross-section which is only of academic interest). Based on this study, of all the blade designs, it has been found that the curve-stacked airfoils exhibit better ‘Rub-tolerant’ behavior. Both experimental and simulation results have predominantly proven the fact that adding curvature to a straight stacked blade through curve-stacked or bow result in reducing the rub induced vibration. While sweep and bow provide some aerodynamic advantages, they are not much helpful in containing the vibrations to a sustainable extent.