Halt, D. W.; Harris, W. L.
1982-01-01
The results reported here are based on applying the method of parametric differentiation (MPD) to transform the nonlinear differential equation governing small-disturbance transonic flow to a linear equation. Implicit approximate factorization and monotone methods were used to accelerate convergence of the linear problem by an order of magnitude over successive line over-relaxation. The relative merits of using MPD are discussed in comparison to conventional small-disturbance applications. Several MPD analyses are performed on an array of airfoils. A design procedure utilizing MPD is discussed and demonstrated for two nonlifting cases.
Shahrokhi, Ava; Jahangirian, Alireza
2010-06-01
A multi-layer perceptron neural network (NN) method is used for efficient estimation of the expensive objective functions in the evolutionary optimization with the genetic algorithm (GA). The estimation capability of the NN is improved by dynamic retraining using the data from successive generations. In addition, the normal distribution of the training data variables is used to determine well-trained parts of the design space for the NN approximation. The efficiency of the method is demonstrated by two transonic airfoil design problems considering inviscid and viscous flow solvers. Results are compared with those of the simple GA and an alternative surrogate method. The total number of flow solver calls is reduced by about 40% using this fitness approximation technique, which in turn reduces the total computational time without influencing the convergence rate of the optimization algorithm. The accuracy of the NN estimation is considerably improved using the normal distribution approach compared with the alternative method.
Design and construction of 2 transonic airfoil models for tests in the NASA Langley C.3-M TCT
Schaechterle, G.; Ludewig, K. H.; Stanewsky, E.; Ray, E. J.
1982-01-01
As part of a NASA/DFVLR cooperation program two transonic airfoils were tested in the NASA Langley 0.3-m TCT. Model design and construction was carried out by DFVLR. The models designed and constructed performed extremely well under cryogenic conditions. Essentially no permanent changes in surface quality and geometric dimensions occurred during the tests. The aerodynamic results from the TCT tests which demonstrate the large sensitivity of the airfoil CAST 10-Z/DOAZ to Reynolds number changes compared well with results from other facilities at ambient temperatures.
ANALYSIS OF TRANSONIC FLOW PAST CUSPED AIRFOILS
Directory of Open Access Journals (Sweden)
Jiří Stodůlka
2015-06-01
Full Text Available Transonic flow past two cusped airfoils is numerically solved and achieved results are analyzed by means of flow behavior and oblique shocks formation.Regions around sharp trailing edges are studied in detail and parameters of shock waves are solved and compared using classical shock polar approach and verified by reduction parameters for symmetric configurations.
Numerical simulation of viscous transonic airfoil flows
Coakley, Thomas J.
1987-01-01
Numerical simulations of transonic airfoil flows using the Reynolds-averaged Navier-Stokes equations and various turbulence models are presented and compared with experimental data. Three different airfoils were investigated under varying flow conditions ranging from subcritical unseparated flows to supercritical separated flows. The turbulence models investigated consisted of three zero-equation models and one two-equation model. For unseparated flows involving weak viscous-inviscid interactions, the four models were comparable in their agreement with experiment. For separated flows involving strong viscous-inviscid interactions, the nonequilibrium zero-equation model of Johnson and King gave the best overall agreement with experiment.
FLEET Velocimetry Measurements on a Transonic Airfoil
Burns, Ross A.; Danehy, Paul M.
2017-01-01
Femtosecond laser electronic excitation tagging (FLEET) velocimetry was used to study the flowfield around a symmetric, transonic airfoil in the NASA Langley 0.3-m TCT facility. A nominal Mach number of 0.85 was investigated with a total pressure of 125 kPa and total temperature of 280 K. Two-components of velocity were measured along vertical profiles at different locations above, below, and aft of the airfoil at angles of attack of 0 deg, 3.5 deg, and 7deg. Measurements were assessed for their accuracy, precision, dynamic range, spatial resolution, and overall measurement uncertainty in the context of the applied flowfield. Measurement precisions as low as 1 m/s were observed, while overall uncertainties ranged from 4 to 5 percent. Velocity profiles within the wake showed sufficient accuracy, precision, and sensitivity to resolve both the mean and fluctuating velocities and general flow physics such as shear layer growth. Evidence of flow separation is found at high angles of attack.
An experimental investigation of a supercritical airfoil at transonic speeds
Mateer, George C.; Seegmiller, H. Lee; Coakley, Thomas J.; Hand, Lawrence A.; Szodruch, Joachim
1987-01-01
Benchmark experimental data obtained in the two-dimensional, transonic flow field surrounding a supercritical airfoil are presented. Airfoil surface and tunnel wall pressure and LDV measurements are used to describe the flow on the model, above the wing and in the wake. Comparisons are made with calculations using the Reynolds-averaged Navier-Stokes equations. The results illustrate the performance of two turbulence models in both separated and attached flows. The largest differences between theory and experiment occurred in separated flows with the Johnson and King turbulence model providing the best estimates.
Fast Euler solver for transonic airfoils. I - Theory. II - Applications
Dadone, Andrea; Moretti, Gino
1988-01-01
Equations written in terms of generalized Riemann variables are presently integrated by inverting six bidiagonal matrices and two tridiagonal matrices, using an implicit Euler solver that is based on the lambda-formulation. The solution is found on a C-grid whose boundaries are very close to the airfoil. The fast solver is then applied to the computation of several flowfields on a NACA 0012 airfoil at various Mach number and alpha values, yielding results that are primarily concerned with transonic flows. The effects of grid fineness and boundary distances are analyzed; the code is found to be robust and accurate, as well as fast.
Transonic aerodynamic design experience
Bonner, E.
1989-01-01
Advancements have occurred in transonic numerical simulation that place aerodynamic performance design into a relatively well developed status. Efficient broad band operating characteristics can be reliably developed at the conceptual design level. Recent aeroelastic and separated flow simulation results indicate that systematic consideration of an increased range of design problems appears promising. This emerging capability addresses static and dynamic structural/aerodynamic coupling and nonlinearities associated with viscous dominated flows.
Transonic buffet control research with two types of shock control bump based on RAE2822 airfoil
Directory of Open Access Journals (Sweden)
Yun TIAN
2017-10-01
Full Text Available Current research shows that the traditional shock control bump (SCB can weaken the intensity of shock and better the transonic buffet performance. The author finds that when SCB is placed downstream of the shock, it can decrease the adverse pressure gradient. This may prevent the shock foot separation bubble to merge with the trailing edge separation and finally improve the buffet performance. Based on RAE2822 airfoil, two types of SCB are designed according to the two different mechanisms. By using Reynolds-averaged Navier-Stokes (RANS and unsteady Reynolds-averaged Navier-Stokes (URANS methods to analyze the properties of RAE2822 airfoil with and without SCB, the results show that the downstream SCB can better the buffet performance under a wide range of freestream Mach number and the steady aerodynamics characteristic is similar to that of RAE2822 airfoil. The traditional SCB can only weaken the intensity of the shock under the design condition. Under the off-design conditions, the SCB does not do much to or even worsen the buffet performance. Indeed, the use of backward bump can flatten the leeward side of the airfoil, and this is similar to the mechanism that supercritical airfoil can weaken the recompression of shock wave.
A finite element method for the computation of transonic flow past airfoils
Eberle, A.
1980-01-01
A finite element method for the computation of the transonic flow with shocks past airfoils is presented using the artificial viscosity concept for the local supersonic regime. Generally, the classic element types do not meet the accuracy requirements of advanced numerical aerodynamics requiring special attention to the choice of an appropriate element. A series of computed pressure distributions exhibits the usefulness of the method.
Intermittent Flow Regimes in a Transonic Fan Airfoil Cascade
Directory of Open Access Journals (Sweden)
J. Lepicovsky
2004-01-01
velocity.To date, this flow behavior has only been observed in a linear transonic cascade. Further research is necessary to confirm this phenomenon occurs in actual transonic fans and is not the by-product of an endwall restricted linear cascade.
Theory of viscous transonic flow over airfoils at high Reynolds number
Melnik, R. E.; Chow, R.; Mead, H. R.
1977-01-01
This paper considers viscous flows with unseparated turbulent boundary layers over two-dimensional airfoils at transonic speeds. Conventional theoretical methods are based on boundary layer formulations which do not account for the effect of the curved wake and static pressure variations across the boundary layer in the trailing edge region. In this investigation an extended viscous theory is developed that accounts for both effects. The theory is based on a rational analysis of the strong turbulent interaction at airfoil trailing edges. The method of matched asymptotic expansions is employed to develop formal series solutions of the full Reynolds equations in the limit of Reynolds numbers tending to infinity. Procedures are developed for combining the local trailing edge solution with numerical methods for solving the full potential flow and boundary layer equations. Theoretical results indicate that conventional boundary layer methods account for only about 50% of the viscous effect on lift, the remaining contribution arising from wake curvature and normal pressure gradient effects.
International Nuclear Information System (INIS)
Ragni, D; Ashok, A; Van Oudheusden, B W; Scarano, F
2009-01-01
The present investigation assesses a procedure to extract the aerodynamic loads and pressure distribution on an airfoil in the transonic flow regime from particle image velocimetry (PIV) measurements. The wind tunnel model is a two-dimensional NACA-0012 airfoil, and the PIV velocity data are used to evaluate pressure fields, whereas lift and drag coefficients are inferred from the evaluation of momentum contour and wake integrals. The PIV-based results are compared to those derived from conventional loads determination procedures involving surface pressure transducers and a wake rake. The method applied in this investigation is an extension to the compressible flow regime of that considered by van Oudheusden et al (2006 Non-intrusive load characterization of an airfoil using PIV Exp. Fluids 40 988–92) at low speed conditions. The application of a high-speed imaging system allows the acquisition in relatively short time of a sufficient ensemble size to compute converged velocity statistics, further translated in turbulent fluctuations included in the pressure and loads calculation, notwithstanding their verified negligible influence in the computation. Measurements are performed at varying spatial resolution to optimize the loads determination in the wake region and around the airfoil, further allowing us to assess the influence of spatial resolution in the proposed procedure. Specific interest is given to the comparisons between the PIV-based method and the conventional procedures for determining the pressure coefficient on the surface, the drag and lift coefficients at different angles of attack. Results are presented for the experiments at a free-stream Mach number M = 0.6, with the angle of attack ranging from 0° to 8°
Ladson, Charles L.; Hill, Acquilla S.; Johnson, William G., Jr.
1987-01-01
Tests were conducted in the 2-D test section of the Langley 0.3-meter Transonic Cryogenic Tunnel on a NACA 0012 airfoil to obtain aerodynamic data as a part of the Advanced Technology Airfoil Test (ATAT) program. The test program covered a Mach number range of 0.30 to 0.82 and a Reynolds number range of 3.0 to 45.0 x 10 to the 6th power. The stagnation pressure was varied between 1.2 and 6.0 atmospheres and the stagnation temperature was varied between 300 K and 90 K to obtain these test conditions. Tabulated pressure distributions and integrated force and moment coefficients are presented as well as plots of the surface pressure distributions. The data are presented uncorrected for wall interference effects and without analysis.
Airfoil, platform, and cooling passage measurements on a rotating transonic high-pressure turbine
Nickol, Jeremy B.
An experiment was performed at The Ohio State University Gas Turbine Laboratory for a film-cooled high-pressure turbine stage operating at design-corrected conditions, with variable rotor and aft purge cooling flow rates. Several distinct experimental programs are combined into one experiment and their results are presented. Pressure and temperature measurements in the internal cooling passages that feed the airfoil film cooling are used as boundary conditions in a model that calculates cooling flow rates and blowing ratio out of each individual film cooling hole. The cooling holes on the suction side choke at even the lowest levels of film cooling, ejecting more than twice the coolant as the holes on the pressure side. However, the blowing ratios are very close due to the freestream massflux on the suction side also being almost twice as great. The highest local blowing ratios actually occur close to the airfoil stagnation point as a result of the low freestream massflux conditions. The choking of suction side cooling holes also results in the majority of any additional coolant added to the blade flowing out through the leading edge and pressure side rows. A second focus of this dissertation is the heat transfer on the rotor airfoil, which features uncooled blades and blades with three different shapes of film cooling hole: cylindrical, diffusing fan shape, and a new advanced shape. Shaped cooling holes have previously shown immense promise on simpler geometries, but experimental results for a rotating turbine have not previously been published in the open literature. Significant improvement from the uncooled case is observed for all shapes of cooling holes, but the improvement from the round to more advanced shapes is seen to be relatively minor. The reduction in relative effectiveness is likely due to the engine-representative secondary flow field interfering with the cooling flow mechanics in the freestream, and may also be caused by shocks and other
Optimization design of airfoil profiles based on the noise of wind turbines
DEFF Research Database (Denmark)
Cheng, Jiangtao; Chen, Jin; Cheng, Jiangtao
2012-01-01
Based on design theory of airfoil profiles and airfoil self-noise prediction model, a new method with the target of the airfoil average efficiency-noise ratio of design ranges for angle of attack had been developed for designing wind turbine airfoils. The airfoil design method was optimized for a...
Integration of Airfoil Design during the design of new blades
Energy Technology Data Exchange (ETDEWEB)
Sartori, L.; Bottasso, L.; Croce, A. [Politecnico di Milano, Milan (Italy); Grasso, F. [ECN Wind Power, Petten (Netherlands)
2013-09-15
Despite the fact that the design of a new blade is a multidisciplinary task, often the different disciplines are combined together at later stage. Looking at the aerodynamic design, it is common practice design/select the airfoils first and then design the blade in terms of chord and twist based on the initial selection of the airfoils. Although this approach is quite diffused, it limits the potentialities of obtaining optimal performance. The present work is focused on investigating the benefits of designing the external shape of the blade including the airfoil shapes together with chord and twist. To accomplish this, a design approach has been developed, where an advanced gradient based optimization algorithm is able to control the shape of the blade. The airfoils described in the work are the NACA 4 digits, while the chord distribution and the twist distribution are described through Bezier curves. In this way, the complexity of the problem is limited while a versatile geometrical description is kept. After the details of the optimization scheme are illustrated, several numerical examples are shown, demonstrating the advantages in terms of performance and development time of integrating the design of the airfoils during the optimization of the blade.
Hess, Robert W.; Seidel, David A.; Igoe, William B.; Lawing, Pierce L.
1987-01-01
Steady and unsteady pressures were measured on a 2-D supercritical airfoil in the Langley Research Center 0.3-m Transonic Cryogenic Tunnel at Reynolds numbers from 6 x 1,000,000 to 35 x 1,000,000. The airfoil was oscillated in pitch at amplitudes from plus or minus .25 degrees to plus or minus 1.0 degrees at frequencies from 5 Hz to 60 Hz. The special requirements of testing an unsteady pressure model in a pressurized cryogenic tunnel are discussed. Selected steady measured data are presented and are compared with GRUMFOIL calculations at Reynolds number of 6 x 1,000,000 and 30 x 1,000,000. Experimental unsteady results at Reynolds numbers of 6 x 1,000,000 and 30 x 1,000,000 are examined for Reynolds number effects. Measured unsteady results at two mean angles of attack at a Reynolds number of 30 x 1,000,000 are also examined.
Optimization design of airfoil profiles based on the noise of wind turbines
DEFF Research Database (Denmark)
Cheng, Jiangtao; Chen, Jin; Cheng, Jiangtao
2012-01-01
Based on design theory of airfoil profiles and airfoil self-noise prediction model, a new method with the target of the airfoil average efficiency-noise ratio of design ranges for angle of attack had been developed for designing wind turbine airfoils. The airfoil design method was optimized...... for a relative thickness of 21% and a new airfoil was obtained. To illustrate the optimization method, the aerodynamic characteristics and noise of the optimized airfoil were calculated and analyzed. Through performance comparison of a DU93-W-210 airfoil and a FFA-W3-211 airfoil which are widely used in wind...
Airfoil family design for large offshore wind turbine blades
International Nuclear Information System (INIS)
Méndez, B; Munduate, X; Miguel, U San
2014-01-01
Wind turbine blades size has scaled-up during last years due to wind turbine platform increase especially for offshore applications. The EOLIA project 2007-2010 (Spanish Goverment funded project) was focused on the design of large offshore wind turbines for deep waters. The project was managed by ACCIONA Energia and the wind turbine technology was designed by ACCIONA Windpower. The project included the design of a wind turbine airfoil family especially conceived for large offshore wind turbine blades, in the order of 5MW machine. Large offshore wind turbines suffer high extreme loads due to their size, in addition the lack of noise restrictions allow higher tip speeds. Consequently, the airfoils presented in this work are designed for high Reynolds numbers with the main goal of reducing blade loads and mantainig power production. The new airfoil family was designed in collaboration with CENER (Spanish National Renewable Energy Centre). The airfoil family was designed using a evolutionary algorithm based optimization tool with different objectives, both aerodynamic and structural, coupled with an airfoil geometry generation tool. Force coefficients of the designed airfoil were obtained using the panel code XFOIL in which the boundary layer/inviscid flow coupling is ineracted via surface transpiration model. The desing methodology includes a novel technique to define the objective functions based on normalizing the functions using weight parameters created from data of airfoils used as reference. Four airfoils have been designed, here three of them will be presented, with relative thickness of 18%, 21%, 25%, which have been verified with the in-house CFD code, Wind Multi Block WMB, and later validated with wind tunnel experiments. Some of the objectives for the designed airfoils concern the aerodynamic behavior (high efficiency and lift, high tangential coefficient, insensitivity to rough conditions, etc.), others concern the geometry (good for structural design
Finite Element Analysis of Transonic Flows over Thin Airfoils. Volume 2. Program User’s Manual
1976-05-01
RKSC, ANDG. XHP, XTE DTM. TFIN, TO, ZTEST, ALPHA, CIRCO (Problem Parameters) IRES « -1. Set T,, Tj, T3 (Start the Integration) Yes Read Inout...steady position. CIRCO Circulation strength for airfoil at its mean position. SQMAC Square of the freestream Mach number. IRES Number of time
Chaparro, Daniel; Fujiwara, Gustavo E. C.; Ting, Eric; Nguyen, Nhan
2016-01-01
The need to rapidly scan large design spaces during conceptual design calls for computationally inexpensive tools such as the vortex lattice method (VLM). Although some VLM tools, such as Vorview have been extended to model fully-supersonic flow, VLM solutions are typically limited to inviscid, subcritical flow regimes. Many transport aircraft operate at transonic speeds, which limits the applicability of VLM for such applications. This paper presents a novel approach to correct three-dimensional VLM through coupling of two-dimensional transonic small disturbance (TSD) solutions along the span of an aircraft wing in order to accurately predict transonic aerodynamic loading and wave drag for transport aircraft. The approach is extended to predict flow separation and capture the attenuation of aerodynamic forces due to boundary layer viscosity by coupling the TSD solver with an integral boundary layer (IBL) model. The modeling framework is applied to the NASA General Transport Model (GTM) integrated with a novel control surface known as the Variable Camber Continuous Trailing Edge Flap (VCCTEF).
Design of wind turbine airfoils based on maximum power coefficient
DEFF Research Database (Denmark)
Cheng, Jiangtao; Chen, Jin; Cheng, Jiangtao
2010-01-01
Based on the blade element momentum (BEM) theory, the power coefficient of a wind turbine can be expressed in function of local tip speed ratio and lift-drag ratio. By taking the power coefficient in a predefined range of angle of attack as the final design objective and combining with an airfoil...... behaviors, noise emission as well as wind turbine service life. To show the performance of the new design technique, a new airfoil with relative thickness of 18% is designed. Comparisons with a wind turbine airfoil (NACA 63418) at Re=2×106 and Re=6×106 for free and fixed transitions show that the new...... airfoil has a higher power efficiency, better designed lift at off-design condition, better stall behavior, less sensitivity to leading edge roughness and lower noise emission. © 2010 Journal of Mechanical Engineering....
Wind turbine airfoil design method with low noise and experimental analysis
DEFF Research Database (Denmark)
Wang, Quan; Chen, Jin; Cheng, Jiangtao
2015-01-01
, though there is a certain difference between the theory results and experiment data. Compared with NACA-64-618 airfoil, the CQU-DTU-B18 airfoil exhibits lower noise, which validates the feasibility of this design method. It is a guide to design wind turbine airfoil with lower noise and to reduce airfoil...
Optimization Criteria and Sailplane Airfoil Design
Czech Academy of Sciences Publication Activity Database
Popelka, Lukáš; Matějka, Milan
2007-01-01
Roč. 30, č. 3 (2007), s. 74-78 ISSN 0744-8996 R&D Projects: GA AV ČR IAA2076403; GA AV ČR(CZ) IAA200760614 Institutional research plan: CEZ:AV0Z20760514 Keywords : aerodynamic optimization * airfoil Subject RIV: BK - Fluid Dynamics
DEFF Research Database (Denmark)
Chen, Jin; Cheng, Jiangtao; Shen, Wenzhong
2013-01-01
Aerodynamic of airfoil performance is closely related to the continuity of its surface curvature, and airfoil profiles with a better aerodynamic performance plays an important role in the design of wind turbine. The surface curvature distribution along the chord direction and pressure distributio...
Design of the new Risoe-A1 airfoil family for wind turbines
Energy Technology Data Exchange (ETDEWEB)
Fuglsang, P.; Dahl, K.S. [Risoe National Lab., Wind Energy and Atmospheric Physics Dept., Roskilde (Denmark)
1999-03-01
A new airfoil family for wind turbines was developed by use of a design method using numerical optimization and the flow solver, XFOIL. The results were evaluated with the Navier-Stokes solver EllipSys2D. The airfoil family constitutes 6 airfoils ranging in thickness from 15% to 30%. The airfoils were designed to have a maximum lift coefficient around 1.5 in natural conditions and high lift-drag ratios below maximum lift. Insensitivity to leading edge roughness was obtained by securing that transition from laminar to turbulent flow on the suction side occurred close to the leading edge just before stall. The airfoil family was designed for a 600 kW wind turbine and provides a basis for further enhancing the characteristics of airfoils for wind turbines and to tailor airfoils for specific rotor sizes and power regulation principles. (au) EFP-95; EFP-98. 16 refs.
Skin design studies for variable camber morphing airfoils
International Nuclear Information System (INIS)
Gandhi, Farhan; Anusonti-Inthra, Phuriwat
2008-01-01
This paper identifies the desirable attributes of a flexible skin of a morphing wing. The study is conducted using airfoil camber morphing as an example. The ideal flex-skin would be highly anisotropic, having a low in-plane axial stiffness but a high out-of-plane flexural stiffness. Reduced skin axial stiffness allows morphing at low actuation cost. However, for some substructure and actuation designs, a lower limit on the skin's in-plane axial stiffness may be required to prevent unacceptable global camber deformation under aerodynamic loads. High flexural stiffness prevents local deformation of skin sections between supports due to aerodynamic pressure loads, and avoids buckling of skin sections under compression as the airfoil cambers under actuation force. For the camber morphing application the strain levels in the flex-skin are not expected to exceed around 2%. If the axial stiffness of the flex-skin is reduced significantly, it may be necessary to consider aerodynamic stiffness (negligible vis-à-vis structural stiffness for classical airfoils) to accurately calculate deformation under loading. The approach followed in the study can be used to identify specifications for the skin and then reverse engineer and design highly anisotropic composite skins that meet the specifications
Airfoil optimization for morphing aircraft
Namgoong, Howoong
Continuous variation of the aircraft wing shape to improve aerodynamic performance over a wide range of flight conditions is one of the objectives of morphing aircraft design efforts. This is being pursued because of the development of new materials and actuation systems that might allow this shape change. The main purpose of this research is to establish appropriate problem formulations and optimization strategies to design an airfoil for morphing aircraft that include the energy required for shape change. A morphing aircraft can deform its wing shape, so the aircraft wing has different optimum shapes as the flight condition changes. The actuation energy needed for moving the airfoil surface is modeled and used as another design objective. Several multi-objective approaches are applied to a low-speed, incompressible flow problem and to a problem involving low-speed and transonic flow. The resulting solutions provide the best tradeoff between low drag, high energy and higher drag, low energy sets of airfoil shapes. From this range of solutions, design decisions can be made about how much energy is needed to achieve a desired aerodynamic performance. Additionally, an approach to model aerodynamic work, which would be more realistic and may allow using pressure on the airfoil to assist a morphing shape change, was formulated and used as part of the energy objective. These results suggest that it may be possible to design a morphing airfoil that exploits the airflow to reduce actuator energy.
Thick airfoil designs for the root of the 10MW INNWIND.EU wind turbine
Mu≁oz, A.; Méndez, B.; Munduate, X.
2016-09-01
The main objective of the “INNWIND.EU” project is to investigate and demonstrate innovative designs for 10-20MW offshore wind turbines and their key components, such as lightweight rotors. In this context, the present paper describes the development of two new airfoils for the blade root region. From the structural point of view, the root is the region in charge of transmitting all the loads of the blade to the hub. Thus, it is very important to include airfoils with adequate structural properties in this region. The present article makes use of high-thickness and blunt trailing edge airfoils to improve the structural characteristics of the airfoils used to build this blade region. CENER's (National Renewable Energy Center of Spain) airfoil design tool uses the airfoil software XFOIL to compute the aerodynamic characteristics of the designed airfoils. That software is based on panel methods which show some problems with the calculation of airfoils with thickness bigger than 35% and with blunt trailing edge. This drawback has been overcome with the development of an empirical correction for XFOIL lift and drag prediction based on airfoil experiments. From the aerodynamic point of view, thick airfoils are known to be very sensitive to surface contamination or turbulent inflow conditions. Consequently, the design optimization takes into account the aerodynamic torque in both clean and contaminated conditions. Two airfoils have been designed aiming to improve the structural and the aerodynamic behaviour of the blade in clean and contaminated conditions. This improvement has been corroborated with Blade Element Momentum (BEM) computations.
Directory of Open Access Journals (Sweden)
M. A. Aziz
2014-01-01
Full Text Available The design of a transonic-fan rotor is optimized using numerical computations of the full three-dimensional Navier-Stokes equations. The CFDRC-ACE multiphysics module, which is a pressure-based solver, is used for the numerical simulation. The code is coupled with simplex optimization algorithm. The optimization process is started from a suitable design point obtained using low fidelity analytical methods that is based on experimental correlations for the pressure losses and blade deviation angle. The fan blade shape is defined by its stacking line and airfoil shape which are considered the optimization parameters. The stacking line is defined by lean, sweep, and skews, while blade airfoil shape is modified considering the thickness and camber distributions. The optimization has been performed to maximize the rotor total pressure ratio while keeping the rotor efficiency and surge margin above certain required values. The results obtained are verified with the experimental data of Rotor 67. In addition, the results of the optimized fan indicate that the optimum design is found to be leaned in the direction of rotation and has a forward sweep from the hub to mean section and backward sweep to the tip. The pressure ratio increases from 1.427 to 1.627 at the design speed and mass flow rate.
Computational design and analysis of flatback airfoil wind tunnel experiment.
Energy Technology Data Exchange (ETDEWEB)
Mayda, Edward A. (University of California, Davis, CA); van Dam, C.P. (University of California, Davis, CA); Chao, David D. (University of California, Davis, CA); Berg, Dale E.
2008-03-01
A computational fluid dynamics study of thick wind turbine section shapes in the test section of the UC Davis wind tunnel at a chord Reynolds number of one million is presented. The goals of this study are to validate standard wind tunnel wall corrections for high solid blockage conditions and to reaffirm the favorable effect of a blunt trailing edge or flatback on the performance characteristics of a representative thick airfoil shape prior to building the wind tunnel models and conducting the experiment. The numerical simulations prove the standard wind tunnel corrections to be largely valid for the proposed test of 40% maximum thickness to chord ratio airfoils at a solid blockage ratio of 10%. Comparison of the computed lift characteristics of a sharp trailing edge baseline airfoil and derived flatback airfoils reaffirms the earlier observed trend of reduced sensitivity to surface contamination with increasing trailing edge thickness.
Design and construction of an airfoil with controlled flap
Amin, Md. Ruhul; Rahman, S. M. Mahbobur; Mashud, Mohammad; Rabbi, Md. Fazle
2017-06-01
For modern aircrafts maneuvering control and reduction of power loss is a matter of great concern in Aerodynamics. Separation of airflow over the wings of aircraft at high angle of attack or at other situations is a hindrance to proper maneuvering control. As flow separation increases drag force on the aircraft, it consumes excess power. For these reasons much effort and research has gone into the design of aerodynamic surfaces which delay flow separation and keep the local flow attached for as long as possible. One of the simple and cost-effective way is to use a hinged flap on the wing of the aircraft, which lifts and self-adjusts to a position dependent on the aerodynamic forces and flap weight due to reversed flow at increasing angle of attack. There is a limitation of this kind of process. At very high angles of attack, the reversed flow would cause the flap to tip forwards entirely and the effect of the flap would vanish. For recovering this limitation an idea of controlling the movement or rotation of the flap has been proposed in this paper. A light surface was selected as a flap and was coupled to the shaft of a servo motor, which was placed on a model airfoil. For controlling the angle of rotation of the motor as well as the flap arbitrarily, an electronic circuit comprising necessary components was designed and applied to the servo motor successfully.
Bionic Design of Wind Turbine Blade Based on Long-Eared Owl’s Airfoil
Directory of Open Access Journals (Sweden)
Weijun Tian
2017-01-01
Full Text Available The main purpose of this paper is to demonstrate a bionic design for the airfoil of wind turbines inspired by the morphology of Long-eared Owl’s wings. Glauert Model was adopted to design the standard blade and the bionic blade, respectively. Numerical analysis method was utilized to study the aerodynamic characteristics of the airfoils as well as the blades. Results show that the bionic airfoil inspired by the airfoil at the 50% aspect ratio of the Long-eared Owl’s wing gives rise to a superior lift coefficient and stalling performance and thus can be beneficial to improving the performance of the wind turbine blade. Also, the efficiency of the bionic blade in wind turbine blades tests increases by 12% or above (up to 44% compared to that of the standard blade. The reason lies in the bigger pressure difference between the upper and lower surface which can provide stronger lift.
High-Tip-Speed, Low-Loading Transonic Fan Stage. Part 1: Aerodynamic and Mechanical Design
Wright, L. C.; Vitale, N. G.; Ware, T. C.; Erwin, J. R.
1973-01-01
A high-tip-speed, low-loading transonic fan stage was designed to deliver an overall pressure ratio of 1.5 with an adiabatic efficiency of 86 percent. The design flow per unit annulus area is 42.0 pounds per square foot. The fan features a hub/tip ratio of 0.46, a tip diameter of 28.74 in. and operates at a design tip speed of 1600 fps. For these design conditions, the rotor blade tip region operates with supersonic inlet and supersonic discharge relative velocities. A sophisticated quasi-three-dimensional characteristic section design procedure was used for the all-supersonic sections and the inlet of the midspan transonic sections. For regions where the relative outlet velocities are supersonic, the blade operates with weak oblique shocks only.
Automated Design and Evaluation of Airfoils for Rotorcraft Applications
2017-12-01
5.8 Performance table pitching moment coecient, Re = 6◊ 106 . . . . 65 5.9 SC1095 lower surface trailing- edge Re” 2 , Re = 6◊ 106 . . . . . . . . 66...b) C-grid (c) trailing- edge Figure 3.4: O-grid around SC1095 airfoil 33 3.3 Performance Table Generation A Python framework named overPy was...framework is developed to automate the generation of airfoil performance tables using the RANS CFD solver OVERFLOW 2.2i allowing the optimization
Design of a new urban wind turbine airfoil using a pressure-load inverse method
Energy Technology Data Exchange (ETDEWEB)
Henriques, J.C.C.; Gato, L.M.C. [IDMEC, Instituto Superior Tecnico, Universidade Tecnica de Lisboa, Av. Rovisco Pais, 1049-001 Lisboa (Portugal); Marques da Silva, F. [LNEC - Laboratorio Nacional de Engenharia Civil, Av. Brasil, 101, 1700-066 Lisboa (Portugal); Estanqueiro, A.I. [INETI - Instituto Nacional de Engenharia, Tecnologia e Inovacao Estrada do Paco do Lumiar, 1649-038 Lisboa (Portugal)
2009-12-15
This paper presents the design methodology of a new wind turbine airfoil that achieves high performance in urban environment by increasing the maximum lift. For this purpose, an inverse method was applied to obtain a new wind turbine blade section with constant pressure-load along the chord, at the design inlet angle. In comparison with conventional blade section designs, the new airfoil has increased maximum lift, reduced leading edge suction peak and controlled soft-stall behaviour, due to a reduction of the adverse pressure gradient on the suction side. Wind tunnel experimental results confirmed the computational results. (author)
DEFF Research Database (Denmark)
Chougle, Prasad Devendra
. Mainly, there is the horizontal axis wind turbine (HAWT) and vertical axis wind turbine (VAWT). HAWTs are more popular than VAWTs due to failure of VAWT commercialization during the late of 1980s on a large scale. However, in recent research work it has been documented that VAWTs are more economical...... for validation. In this PhD research, a development of wind turbine rotor is planned based on the multi-element airfoil technology used in aviation for aeroplanes. A method of experimental and numerical analysis is combined together for successful research. A double-element airfoil design is carried out...
Computational Design and Analysis of a Transonic Natural Laminar Flow Wing for a Wind Tunnel Model
Lynde, Michelle N.; Campbell, Richard L.
2017-01-01
A natural laminar flow (NLF) wind tunnel model has been designed and analyzed for a wind tunnel test in the National Transonic Facility (NTF) at the NASA Langley Research Center. The NLF design method is built into the CDISC design module and uses a Navier-Stokes flow solver, a boundary layer profile solver, and stability analysis and transition prediction software. The NLF design method alters the pressure distribution to support laminar flow on the upper surface of wings with high sweep and flight Reynolds numbers. The method addresses transition due to attachment line contamination/transition, Gortler vortices, and crossflow and Tollmien-Schlichting modal instabilities. The design method is applied to the wing of the Common Research Model (CRM) at transonic flight conditions. Computational analysis predicts significant extents of laminar flow on the wing upper surface, which results in drag savings. A 5.2 percent scale semispan model of the CRM NLF wing will be built and tested in the NTF. This test will aim to validate the NLF design method, as well as characterize the laminar flow testing capabilities in the wind tunnel facility.
Design Of An Aerodynamic Measurement System For Unmanned Aerial Vehicle Airfoils
Directory of Open Access Journals (Sweden)
L. Velázquez-Araque
2012-10-01
Full Text Available This paper presents the design and validation of a measurement system for aerodynamic characteristics of unmanned aerial vehicles. An aerodynamic balance was designed in order to measure the lift, drag forces and pitching moment for different airfoils. During the design process, several aspects were analyzed in order to produce an efficient design, for instance the range of changes of the angle of attack with and a small increment and the versatility of being adapted to different type of airfoils, since it is a wire balance it was aligned and calibrated as well. Wind tunnel tests of a two dimensional NACA four digits family airfoil and four different modifications of this airfoil were performed to validate the aerodynamic measurement system. The modification of this airfoil was made in order to create a blowing outlet with the shape of a step on the suction surface. Therefore, four different locations along the cord line for this blowing outlet were analyzed. This analysis involved the aerodynamic performance which meant obtaining lift, drag and pitching moment coefficients curves as a function of the angle of attack experimentally for the situation where the engine of the aerial vehicle is turned off, called the no blowing condition, by means of wind tunnel tests. The experiments were performed in a closed circuit wind tunnel with an open test section. Finally, results of the wind tunnel tests were compared with numerical results obtained by means of computational fluid dynamics as well as with other experimental references and found to be in good agreement.
Check-Standard Testing Across Multiple Transonic Wind Tunnels with the Modern Design of Experiments
Deloach, Richard
2012-01-01
This paper reports the result of an analysis of wind tunnel data acquired in support of the Facility Analysis Verification & Operational Reliability (FAVOR) project. The analysis uses methods referred to collectively at Langley Research Center as the Modern Design of Experiments (MDOE). These methods quantify the total variance in a sample of wind tunnel data and partition it into explained and unexplained components. The unexplained component is further partitioned in random and systematic components. This analysis was performed on data acquired in similar wind tunnel tests executed in four different U.S. transonic facilities. The measurement environment of each facility was quantified and compared.
Design of a 21 m blade with Risø-A1 airfoils for active stall controlled wind turbines
DEFF Research Database (Denmark)
Fuglsang, Peter; Sangill, O.; Hansen, P.
2002-01-01
This is the final report, from the project, "Design of a Rotor/Airfoil Family for Active Stall-regulated Wind Turbines by Use of Multi-point Optimization". It describes the full scale testing of a 21 m wind turbine blade specially designed for active stallregulation. Design objectives were...... increased ratio of produced energy to turbine loads and more stable power control characteristics. Both were taken directly into account during the design of the blade using numerical optimization. The blade used theRisø-A1 airfoil family, which was specially designed for operation on wind turbine blades...... be concluded that the new LM 21.0 ASR blade could replace the LM 21.0P leading to improved cost efficiency and that the Risø-A1 airfoils were well suited for active stall control. With the newestablished knowledge of the actual airfoil characteristics, a possible future blade design could be made also...
Dynamic Stall Characteristics of Drooped Leading Edge Airfoils
Sankar, Lakshmi N.; Sahin, Mehmet; Gopal, Naveen
2000-01-01
Helicopters in high-speed forward flight usually experience large regions of dynamic stall over the retreating side of the rotor disk. The rapid variations in the lift and pitching moments associated with the stall process can result in vibratory loads, and can cause fatigue and failure of pitch links. In some instances, the large time lag between the aerodynamic forces and the blade motion can trigger stall flutter. A number of techniques for the alleviation of dynamic stall have been proposed and studied by researchers. Passive and active control techniques have both been explored. Passive techniques include the use of high solidity rotors that reduce the lift coefficients of individual blades, leading edge slots and leading edge slats. Active control techniques include steady and unsteady blowing, and dynamically deformable leading edge (DDLE) airfoils. Considerable amount of experimental and numerical data has been collected on the effectiveness of these concepts. One concept that has not received as much attention is the drooped-leading edge airfoil idea. It has been observed in wind tunnel studies and flight tests that drooped leading edge airfoils can have a milder dynamic stall, with a significantly milder load hysteresis. Drooped leading edge airfoils may not, however, be suitable at other conditions, e.g. in hover, or in transonic flow. Work needs to be done on the analysis and design of drooped leading edge airfoils for efficient operation in a variety of flight regimes (hover, dynamic stall, and transonic flow). One concept that is worthy of investigation is the dynamically drooping airfoil, where the leading edge shape is changed roughly once-per-rev to mitigate the dynamic stall.
International Nuclear Information System (INIS)
Lee, Sae Il; Lee, Dong Ho; Kim, Kyu Hong; Park, Tae Choon; Lim, Byeung Jun; Kang, Young Seok
2013-01-01
The multidisciplinary design optimization method, which integrates aerodynamic performance and structural stability, was utilized in the development of a single-stage transonic axial compressor. An approximation model was created using artificial neural network for global optimization within given ranges of variables and several design constraints. The genetic algorithm was used for the exploration of the Pareto front to find the maximum objective function value. The final design was chosen after a second stage gradient-based optimization process to improve the accuracy of the optimization. To validate the design procedure, numerical simulations and compressor tests were carried out to evaluate the aerodynamic performance and safety factor of the optimized compressor. Comparison between numerical optimal results and experimental data are well matched. The optimum shape of the compressor blade is obtained and compared to the baseline design. The proposed optimization framework improves the aerodynamic efficiency and the safety factor.
Application of Artificial Neural Networks to the Design of Turbomachinery Airfoils
Rai, Man Mohan; Madavan, Nateri
1997-01-01
Artificial neural networks are widely used in engineering applications, such as control, pattern recognition, plant modeling and condition monitoring to name just a few. In this seminar we will explore the possibility of applying neural networks to aerodynamic design, in particular, the design of turbomachinery airfoils. The principle idea behind this effort is to represent the design space using a neural network (within some parameter limits), and then to employ an optimization procedure to search this space for a solution that exhibits optimal performance characteristics. Results obtained for design problems in two spatial dimensions will be presented.
Morris, A. L.; Halle, J. E.; Kennedy, E. E.
1972-01-01
A single stage fan with a tip speed of 1800 ft/sec (548.6m/sec) and hub/tip ratio of 0.5 was designed to produce a pressure ratio of 2.285:1 with an adiabatic efficiency of 84.0%. The design flow per inlet annulus area is 38.7 lbm/sq ft-sec (188.9KG/sqm-sec). Rotor blades have modified multiple-circular-arc and precompression airfoil sections. The stator vanes have multiple-circular-arc airfoil sections.
Artificial viscosity method for the design of supercritical airfoils. [Analysis code H
Energy Technology Data Exchange (ETDEWEB)
McFadden, G.B.
1979-07-01
The need for increased efficiency in the use of our energy resources has stimulated applied research in many areas. Recently progress has been made in the field of aerodynamics, where the development of the supercritical wing promises significant savings in the fuel consumption of aircraft operating near the speed of sound. Computational transonic aerodynamics has proved to be a useful tool in the design and evaluation of these wings. A numerical technique for the design of two-dimensional supercritical wing sections with low wave drag is presented. The method is actually a design mode of the analysis code H developed by Bauer, Garabedian, and Korn. This analysis code gives excellent agreement with experimental results and is used widely by the aircraft industry. The addition of a conceptually simple design version should make this code even more useful to the engineering public.
Introduction to transonic aerodynamics
Vos, Roelof
2015-01-01
Written to teach students the nature of transonic flow and its mathematical foundation, this book offers a much-needed introduction to transonic aerodynamics. The authors present a quantitative and qualitative assessment of subsonic, supersonic, and transonic flow around bodies in two and three dimensions. The book reviews the governing equations and explores their applications and limitations as employed in modeling and computational fluid dynamics. Some concepts, such as shock and expansion theory, are examined from a numerical perspective. Others, including shock-boundary-layer interaction, are discussed from a qualitative point of view. The book includes 60 examples and more than 200 practice problems. The authors also offer analytical methods such as Method of Characteristics (MOC) that allow readers to practice with the subject matter. The result is a wealth of insight into transonic flow phenomena and their impact on aircraft design, including compressibility effects, shock and expansion waves, sho...
Design and validation of the high performance and low noise CQU-DTU-LN1 airfoils
DEFF Research Database (Denmark)
Cheng, Jiangtao; Zhu, Wei Jun; Fischer, Andreas
2014-01-01
objective, such as airfoil lift coefficient, drag coefficient and lift-drag ratio, and minimizes trailing edge noise as a constraint. To express airfoil shape, an analytical expression is used. One of the main advantages of the present designmethod is that it produces a highly smooth airfoil shape that can...
Castiñeira, Esther; Solís, Irene; Argüelles, K.M. (Katia); Velarde, Sandra; Fernández, J.M. (Jesús); González, Jose
2016-01-01
In this work, the capability of simple numerical models with coarse grids to predict performance coefficients in wind turbine airfoils is explored. A wide range of simulations were performed for a typical wind turbine profile, under the main criteria of design simplicity and low calculation time. The solutions were computed over different mesh sizes using a two-dimensional Reynolds-Average Navier-Stockes (2D-RANS) approach. Spalart-Allmaras, k-ε and k-omega turbulence models were run in the s...
Yang, Y.; Özgen, S.
2017-06-01
During the last few decades, CFD (Computational Fluid Dynamics) has developed greatly and has become a more reliable tool for the conceptual phase of aircraft design. This tool is generally combined with an optimization algorithm. In the optimization phase, the need for regenerating the computational mesh might become cumbersome, especially when the number of design parameters is high. For this reason, several mesh generation and deformation techniques have been developed in the past decades. One of the most widely used techniques is the Spring Analogy. There are numerous spring analogy related techniques reported in the literature: linear spring analogy, torsional spring analogy, semitorsional spring analogy, and ball vertex spring analogy. This paper gives the explanation of linear spring analogy method and angle inclusion in the spring analogy method. In the latter case, two di¨erent solution methods are proposed. The best feasible method will later be used for two-dimensional (2D) Airfoil Design Optimization with objective function being to minimize sectional drag for a required lift coe©cient at di¨erent speeds. Design variables used in the optimization include camber and thickness distribution of the airfoil. SU2 CFD is chosen as the §ow solver during the optimization procedure. The optimization is done by using Phoenix ModelCenter Optimization Tool.
Unsteady transonic aerodynamics
International Nuclear Information System (INIS)
Nixon, D.
1989-01-01
Various papers on unsteady transonic aerodynamics are presented. The topics addressed include: physical phenomena associated with unsteady transonic flows, basic equations for unsteady transonic flow, practical problems concerning aircraft, basic numerical methods, computational methods for unsteady transonic flows, application of transonic flow analysis to helicopter rotor problems, unsteady aerodynamics for turbomachinery aeroelastic applications, alternative methods for modeling unsteady transonic flows
Supercritical Airfoil Coordinates
National Aeronautics and Space Administration — Rectangular Supercritical Wing (Ricketts) - design and measured locations are provided in an Excel file RSW_airfoil_coordinates_ricketts.xls . One sheet is with Non...
OUT Success Stories: Advanced Airfoils for Wind Turbines
International Nuclear Information System (INIS)
Jones, J.; Green, B.
2000-01-01
New airfoils have substantially increased the aerodynamic efficiency of wind turbines. It is clear that these new airfoils substantially increased energy output from wind turbines. Virtually all new blades built in this country today use these advanced airfoil designs
Wind Tunnel Evaluation of a Model Helicopter Main-Rotor Blade With Slotted Airfoils at the Tip
Noonan, Kevin W.; Yeager, William T., Jr.; Singleton, Jeffrey D.; Wilbur, Matthew L.; Mirick, Paul H.
2001-01-01
Data for rotors using unconventional airfoils are of interest to permit an evaluation of this technology's capability to meet the U.S. Army's need for increased helicopter mission effectiveness and improved safety and survivability. Thus, an experimental investigation was conducted in the Langley Transonic Dynamics Tunnel (TDT) to evaluate the effect of using slotted airfoils in the rotor blade tip region (85 to 100 percent radius) on rotor aerodynamic performance and loads. Four rotor configurations were tested in forward flight at advance ratios from 0.15 to 0.45 and in hover in-ground effect. The hover tip Mach number was 0.627, which is representative of a design point of 4000-ft geometric altitude and a temperature of 95 F. The baseline rotor configuration had a conventional single-element airfoil in the tip region. A second rotor configuration had a forward-slotted airfoil with a -6 deg slat, a third configuration had a forward-slotted airfoil with a -10 slat, and a fourth configuration had an aft-slotted airfoil with a 3 deg flap (trailing edge down). The results of this investigation indicate that the -6 deg slat configuration offers some performance and loads benefits over the other three configurations.
The SNL100-03 Blade: Design Studies with Flatback Airfoils for the Sandia 100-meter Blade.
Energy Technology Data Exchange (ETDEWEB)
Griffith, Daniel; Richards, Phillip William
2014-09-01
A series of design studies were performed to inv estigate the effects of flatback airfoils on blade performance and weight for large blades using the Sandi a 100-meter blade designs as a starting point. As part of the study, the effects of varying the blade slenderness on blade structural performance was investigated. The advantages and disadvantages of blad e slenderness with respect to tip deflection, flap- wise & edge-wise fatigue resistance, panel buckling capacity, flutter speed, manufacturing labor content, blade total weight, and aerodynamic design load magn itude are quantified. Following these design studies, a final blade design (SNL100-03) was prod uced, which was based on a highly slender design using flatback airfoils. The SNL100-03 design with flatback airfoils has weight of 49 tons, which is about 16% decrease from its SNL100-02 predecessor that used conventional sharp trailing edge airfoils. Although not systematically optimized, the SNL100 -03 design study provides an assessment of and insight into the benefits of flatback airfoils for la rge blades as well as insights into the limits or negative consequences of high blade slenderness resulting from a highly slender SNL100-03 planform as was chosen in the final design definition. This docum ent also provides a description of the final SNL100-03 design definition and is intended to be a companion document to the distribution of the NuMAD blade model files for SNL100-03, which are made publicly available. A summary of the major findings of the Sandia 100-meter blade development program, from the initial SNL100-00 baseline blade through the fourth SNL100-03 blade study, is provided. This summary includes the major findings and outcomes of blade d esign studies, pathways to mitigate the identified large blade design drivers, and tool development that were produced over the course of this five-year research program. A summary of large blade tec hnology needs and research opportunities is also presented.
Optimal Design of Airfoil with High Aspect Ratio in Unmanned Aerial Vehicles
Kyoungwoo Park; Ji-Won Han; Hyo-Jae Lim; Byeong-Sam Kim; Juhee Lee
2008-01-01
Shape optimization of the airfoil with high aspect ratio of long endurance unmanned aerial vehicle (UAV) is performed by the multi-objective optimization technology coupled with computational fluid dynamics (CFD). For predicting the aerodynamic characteristics around the airfoil the high-fidelity Navier-Stokes solver is employed and SMOGA (Simple Multi-Objective Genetic Algorithm), which is developed by authors, is used for solving the multi-objective optimization problem...
Wood, M. E.; Newman, D. A.
An acoustic liner has been designed and manufactured for use in the ARA transonic wind tunnel to provide an acoustically acceptable environment for propeller noise testing up to high subsonic Mach number. Details of the aerodynamic design and development are presented and calibration of the liner with propeller model support systems is included. It is shown how the design of the acoustic treatment was aided by the use of a theoretical model for the tunnel reverberant field. An acoustic development program was undertaken involving horn tests to improve the quality of the liner. The success of this is demonstrated by propeller noise results. These results also provided the basis for definition of the practical acoustic test regime of the ARA lined tunnel suitable for the accurate measurement of propeller noise.
Numerical Solution of Compressible Steady Flows around the RAE 2822 Airfoil
Directory of Open Access Journals (Sweden)
Kryštůfek P.
2014-03-01
Full Text Available The article presents results of a numerical solution of subsonic, transonic and supersonic flows described by the system of Navier-Stokes equations in 2D laminar compressible flows around the RAE 2822 airfoil. Authors used FVM multistage Runge-Kutta method to numerically solve the flows around the RAE 2822 airfoil.
Numerical Solution of Compressible Steady Flows around the NACA 0012 Airfoil
Kryštůfek, P.; Kozel, K.
2013-04-01
The article presents results of a numerical solution of subsonic and transonic flows described by the system of Navier-Stokes equations in 2D laminar compressible flows around the NACA 0012 airfoil. Authors used Runge-Kutta method to numerically solve the flows around the NACA 0012 airfoil.
Numerical Solution of Compressible Steady Flows around the NACA 0012 Airfoil
Directory of Open Access Journals (Sweden)
Kozel K
2013-04-01
Full Text Available The article presents results of a numerical solution of subsonic and transonic flows described by the system of Navier-Stokes equations in 2D laminar compressible flows around the NACA 0012 airfoil. Authors used Runge-Kutta method to numerically solve the flows around the NACA 0012 airfoil.
Numerical Solution of Compressible Steady Flows around the RAE 2822 Airfoil
Kryštůfek, P.; Kozel, K.
2014-03-01
The article presents results of a numerical solution of subsonic, transonic and supersonic flows described by the system of Navier-Stokes equations in 2D laminar compressible flows around the RAE 2822 airfoil. Authors used FVM multistage Runge-Kutta method to numerically solve the flows around the RAE 2822 airfoil.
Research on design methods and aerodynamics performance of CQUDTU-B21 airfoil
DEFF Research Database (Denmark)
Chen, Jin; Cheng, Jiangtao; Wen, Zhong Shen
2012-01-01
-DTU-B21 airfoil. To validate the optimization results, the comparison of the aerodynamics performance by XFOIL and wind tunnels test respectively at Re=3×106 is made between the CQU-DTU-B21 and DU93-W-210 which is widely used in wind turbines. © (2012) Trans Tech Publications, Switzerland....
Holography and LDV techniques, their status and use in airfoil research
Johnson, D. A.; Bachalo, W. D.
1978-01-01
The measurement capabilities of laser velocimetry and holographic interferometry in transonic airfoil testing were demonstrated. Presented are representative results obtained with these two nonintrusive techniques on a 15.24 cm chord airfoil section. These results include the density field about the airfoil, flow angles in the inviscid flow and viscous flow properties including the turbulent Reynolds stresses. The accuracies of the density fields obtained by interferometry were verified from comparisons with surface pressure and laser velocimeter measurements.
Design and verification of the Risø-B1 airfoil family for wind turbines
DEFF Research Database (Denmark)
Fuglsang, P.; Bak, C.; Gaunaa, M.
2004-01-01
severe roughness caused reductions in maximum lift between 12% and 27%. Results for the Risø-B1-24 airfoil showed a maximum lift coefficient of 1.62. The standard case leading edge roughness caused a drop in maximum lift of 7.4%. Vortex generators and Gurney flaps in combination could increase maximum...... wind tunnel, Denmark, at a Reynolds number of 1.6x10(6). For both airfoils the predicted target characteristics were met. Results for Risø-B1-18 showed a maximum lift coefficient of 1.64. A standard case of zigzag tape leading edge roughness caused a drop in maximum lift of only 3.7%. Cases of more...
Turbine airfoil to shround attachment
Campbell, Christian X; Morrison, Jay A; James, Allister W; Snider, Raymond G; Eshak, Daniel M; Marra, John J; Wessell, Brian J
2014-05-06
A turbine airfoil (31) with an end portion (42) that tapers (44) toward the end (43) of the airfoil. A ridge (46) extends around the end portion. It has proximal (66) and distal (67) sides. A shroud platform (50) is bi-cast onto the end portion around the ridge without bonding. Cooling shrinks the platform into compression (62) on the end portion (42) of the airfoil. Gaps between the airfoil and platform are formed using a fugitive material (56) in the bi-casting stage. These gaps are designed in combination with the taper angle (44) to accommodate differential thermal expansion while maintaining a gas seal along the contact surfaces. The taper angle (44) may vary from lesser on the pressure side (36) to greater on the suction side (38) of the airfoil. A collar portion (52) of the platform provides sufficient contact area for connection stability.
Airfoil design: Finding the balance between design lift and structural stiffness
DEFF Research Database (Denmark)
Bak, Christian; Gaudern, Nicholas; Zahle, Frederik
2014-01-01
When upscaling wind turbine blades there is an increasing need for high levels of structural efficiency. In this paper the relationships between the aerodynamic characteristics; design lift and lift-drag ratio; and the structural characteristics were investigated. Using a unified optimization setup...
Study on optimal design of wind turbine blade airfoil and its application
International Nuclear Information System (INIS)
Sun, Min Young; Kim, Dong Yong; Lim, Jae Kyoo
2012-01-01
This study was carried out with two goals. One was the development of a model of a wind turbine blade airfoil and the other was the application of the folding blade. In general, in large sized (MW) wind turbines, damage is prevented in small wind turbines since equipment costs and maintenance costs are high, and therefore, the blade will cause serious damage. The wind turbine proposed in this study does not require maintenance, and the blades do not break during high winds because they are folded in accordance with changes in the wind speed. But generators are not cut out, while maintaining a constant angle will continue to produce. The focus of this study, the wind turbine is continued by folding blade system in strong winds and gusts without stopping production
Aerodynamic Analysis Over Double Wedge Airfoil
Prasad, U. S.; Ajay, V. S.; Rajat, R. H.; Samanyu, S.
2017-05-01
Aeronautical studies are being focused more towards supersonic flights and methods to attain a better and safer flight with highest possible performance. Aerodynamic analysis is part of the whole procedure, which includes focusing on airfoil shapes which will permit sustained flight of aircraft at these speeds. Airfoil shapes differ based on the applications, hence the airfoil shapes considered for supersonic speeds are different from the ones considered for Subsonic. The present work is based on the effects of change in physical parameter for the Double wedge airfoil. Mach number range taken is for transonic and supersonic. Physical parameters considered for the Double wedge case with wedge angle (ranging from 5 degree to 15 degree. Available Computational tools are utilized for analysis. Double wedge airfoil is analysed at different Angles of attack (AOA) based on the wedge angle. Analysis is carried out using fluent at standard conditions with specific heat ratio taken as 1.4. Manual calculations for oblique shock properties are calculated with the help of Microsoft excel. MATLAB is used to form a code for obtaining shock angle with Mach number and wedge angle at the given parameters. Results obtained from manual calculations and fluent analysis are cross checked.
Abramova, K. A.; Petrov, A. V.; Potapchick, A. V.; Soudakov, V. G.
2016-10-01
Numerical and experimental investigations of transonic buffet control by tangential jet blowing are presented. To suppress the shock-induced boundary layer separation and the buffet at transonic speeds, compressed air jet is blown through a small slot nozzle tangentially to the upper surface of the supercritical airfoil. Numerical simulations were carried out on the basis of the unsteady Reynolds averaged Navier-Stokes (URANS) equations. Experimental studies of the tangential jet blowing were performed in the transonic wind tunnel T-112 of TsAGI. Results show that the jet moves the shock downstream, increases lift, suppresses flow separation under shock foot and delays buffet onset.
Yates, E. C., Jr.; Wynne, E. C.; Farmer, M. G.; Desmarais, R. N.
1981-01-01
Use of a supercritical airfoil can adversely affect wing flutter speeds in the transonic range. As adequate theories for three dimensional unsteady transonic flow are not yet available, the modified strip analysis was used to predict the transonic flutter boundary for the supercritical wing. The steady state spanwise distributions of section lift curve slope and aerodynamic center, required as input for the flutter calculations, were obtained from pressure distributions. The calculated flutter boundary is in agreement with experiment in the subsonic range. In the transonic range, a transonic bucket is calculated which closely resembles the experimental one with regard to both shape and depth, but it occurs at about 0.04 Mach number lower than the experimental one.
The design and development of an acoustic test section for the ARA transonic wind tunnel
Wood, M. E.; Newman, D. A.
1988-11-01
The structural, aerodynamic and acoustic design of a lined test section necessary to produce a sufficiently reflection-free environment for noise tests in the Aircraft Research Association (ARA) tunnel at Mach numbers up to 0.85 are described. The aerodynamic design was aided by computational fluid dynamics methods and the acoustic design by a ray-theory model. Results of the aerodynamic calibration of the liner and preliminary acoustic and performance tests on two model propellers are presented. Development work carried out to improve the liner acoustically and aerodynamically is described.
Transonic Experimental Research Facility
Federal Laboratory Consortium — The Transonic Experimental Research Facility evaluates aerodynamics and fluid dynamics of projectiles, smart munitions systems, and sub-munitions dispensing systems;...
Hylton, L. D.; Nirmalan, V.; Sultanian, B. K.; Kaufman, R. M.
1987-10-01
The experimental data obtained in this program gives insight into the physical phenomena that occur on a film cooled airfoil, and should provide a relevant data base for verification of new design tools. Results indicate that the downstream film cooling process is a complex function of the thermal dilution and turbulence augmentation parameters with trends actually reversing as blowing strength and coolant-to-gas temperature ratio varied. The pressure surface of the airfoil is shown to exhibit a considerably higher degree of sensitivity to changes in the film cooling parameters and, consequently, should prove to be more of a challenge than the suction surface in accurately predicting heat transfer levels with downsteam film cooling.
Numerical Solution of Inviscid Compressible Steady Flows around the RAE 2822 Airfoil
Directory of Open Access Journals (Sweden)
Kryštůfek P.
2015-01-01
Full Text Available The article presents results of a numerical solution of subsonic, transonic and supersonic flows described by the system of Euler equations in 2D compressible flows around the RAE 2822 airfoil. Authors used FVM multistage Runge-Kutta method to numerically solve the flows around the RAE 2822 airfoil. The results are compared with the solution using the software Ansys Fluent 15.0.7.
Numerical Solution of Inviscid Compressible Steady Flows around the RAE 2822 Airfoil
Kryštůfek, P.; Kozel, K.
2015-05-01
The article presents results of a numerical solution of subsonic, transonic and supersonic flows described by the system of Euler equations in 2D compressible flows around the RAE 2822 airfoil. Authors used FVM multistage Runge-Kutta method to numerically solve the flows around the RAE 2822 airfoil. The results are compared with the solution using the software Ansys Fluent 15.0.7.
Hybrid Optimization for Wind Turbine Thick Airfoils
Energy Technology Data Exchange (ETDEWEB)
Grasso, F. [ECN Wind Energy, Petten (Netherlands)
2012-06-15
One important element in aerodynamic design of wind turbines is the use of specially tailored airfoils to increase the ratio of energy capture and reduce cost of energy. This work is focused on the design of thick airfoils for wind turbines by using numerical optimization. A hybrid scheme is proposed in which genetic and gradient based algorithms are combined together to improve the accuracy and the reliability of the design. Firstly, the requirements and the constraints for this class of airfoils are described; then, the hybrid approach is presented. The final part of this work is dedicated to illustrate a numerical example regarding the design of a new thick airfoil. The results are discussed and compared to existing airfoils.
Energy Technology Data Exchange (ETDEWEB)
G.E. Fuchs
2007-12-31
Recent initiatives for fuel flexibility, increased efficiency and decreased emissions in power generating industrial gas turbines (IGT's), have highlighted the need for the development of techniques to produce large single crystal or columnar grained, directionally solidified Ni-base superalloy turbine blades and vanes. In order to address the technical difficulties of producing large single crystal components, a program has been initiated to, using computational materials science, better understand how alloy composition in potential IGT alloys and solidification conditions during processing, effect castability, defect formation and environmental resistance. This program will help to identify potential routes for the development of high strength, corrosion resistant airfoil/vane alloys, which would be a benefit to all IGT's, including small IGT's and even aerospace gas turbines. During the first year, collaboration with Siemens Power Corporation (SPC), Rolls-Royce, Howmet and Solar Turbines has identified and evaluated about 50 alloy compositions that are of interest for this potential application. In addition, alloy modifications to an existing alloy (CMSX-4) were also evaluated. Collaborating with SPC and using computational software at SPC to evaluate about 50 alloy compositions identified 5 candidate alloys for experimental evaluation. The results obtained from the experimentally determined phase transformation temperatures did not compare well to the calculated values in many cases. The effects of small additions of boundary strengtheners (i.e., C, B and N) to CMSX-4 were also examined. The calculated phase transformation temperatures were somewhat closer to the experimentally determined values than for the 5 candidate alloys, discussed above. The calculated partitioning coefficients were similar for all of the CMSX-4 alloys, similar to the experimentally determined segregation behavior. In general, it appears that computational materials
Advanced Airfoils Boost Helicopter Performance
2007-01-01
Carson Helicopters Inc. licensed the Langley RC4 series of airfoils in 1993 to develop a replacement main rotor blade for their Sikorsky S-61 helicopters. The company's fleet of S-61 helicopters has been rebuilt to include Langley's patented airfoil design, and the helicopters are now able to carry heavier loads and fly faster and farther, and the main rotor blades have twice the previous service life. In aerial firefighting, the performance-boosting airfoils have helped the U.S. Department of Agriculture's Forest Service control the spread of wildfires. In 2003, Carson Helicopters signed a contract with Ducommun AeroStructures Inc., to manufacture the composite blades for Carson Helicopters to sell
Transonic Dynamics Tunnel (TDT)
Federal Laboratory Consortium — The Transonic Dynamics Tunnel (TDT) is a continuous flow wind-tunnel facility capable of speeds up to Mach 1.2 at stagnation pressures up to one atmosphere. The TDT...
Laser velocimetry applied to transonic and supersonic aerodynamics
Johnson, D. A.; Bachalo, W. D.; Moddaress, D.
1976-01-01
As a further demonstration of the capabilities of laser velocity in compressible aerodynamics, measurements obtained in a Mach 2.9 separated turbulent boundary layer and in the transonic flow past a two-dimensional airfoil section are presented and compared to data realized by conventional techniques. In the separated-flow study, the comparisons were made against pitot-static pressure data. Agreement in mean velocities was realized where the pressure measurements could be considered reliable; however, in regions of instantaneous reverse velocities, the laser results were found to be consistent with the physics of the flow whereas the pressure data were not. The laser data obtained in regions of extremely high turbulence suggest that velocity biasing does not occur if the particle occurrence rate is low relative to the turbulent fluctuation rate. Streamwise turbulence intensities are also presented. In the transonic airfoil study, velocity measurements obtained immediately outside the upper surface boundary layer of a 6-inch chord MACA 64A010 airfoil are compared to edge velocities inferred from surface pressure measurements. For free-stream Mach numbers of 0.6 and 0.8, the agreement in results was very good. Dual scatter optical arrangements in conjunction with a single particle, counter-type signal processor were employed in these investigations. Half-micron-diameter polystyrene spheres and naturally occurring condensed oil vapor acted as light scatterers in the two respective flows. Bragg-cell frequency shifting was utilized in the separated flow study.
Nonlinear aeroelastic behavior of compliant airfoils
International Nuclear Information System (INIS)
Thwapiah, G; Campanile, L F
2010-01-01
Since the beginning of aviation and up to the present time, airfoils have always been built as rigid structures. They are designed to fly under their divergence speed in order to avoid static aeroelastic instabilities and the resulting large deformations, which are not compatible with the typically low compliance of such airfoils. In recent years, research on airfoil morphing has generated interest in innovative ideas like the use of compliant systems, i.e. systems built to allow for large deformations without failure, in airfoil construction. Such systems can operate in the neighborhood of divergence and take advantage of large aeroelastic servo-effects. This, in turn, allows compact, advanced actuators to control the airfoil's deformation and loads, and hence complement or even replace conventional flaps. In order to analyze and design such compliant, active aeroelastic structures a nonlinear approach to static aeroelasticity is needed, which takes into account the effect of large deformations on aerodynamics and structure. Such an analytical approach is presented in this paper and applied to a compliant passive airfoil as the preliminary step in the realization of a piezoelectrically driven, active aeroelastic airfoil. Wind tunnel test results are also presented and compared with the analytic prediction. The good agreement and the observed behavior in the wind tunnel give confidence in the potential of this innovative idea
Nonlinear aeroelastic behavior of compliant airfoils
Thwapiah, G.; Campanile, L. F.
2010-03-01
Since the beginning of aviation and up to the present time, airfoils have always been built as rigid structures. They are designed to fly under their divergence speed in order to avoid static aeroelastic instabilities and the resulting large deformations, which are not compatible with the typically low compliance of such airfoils. In recent years, research on airfoil morphing has generated interest in innovative ideas like the use of compliant systems, i.e. systems built to allow for large deformations without failure, in airfoil construction. Such systems can operate in the neighborhood of divergence and take advantage of large aeroelastic servo-effects. This, in turn, allows compact, advanced actuators to control the airfoil's deformation and loads, and hence complement or even replace conventional flaps. In order to analyze and design such compliant, active aeroelastic structures a nonlinear approach to static aeroelasticity is needed, which takes into account the effect of large deformations on aerodynamics and structure. Such an analytical approach is presented in this paper and applied to a compliant passive airfoil as the preliminary step in the realization of a piezoelectrically driven, active aeroelastic airfoil. Wind tunnel test results are also presented and compared with the analytic prediction. The good agreement and the observed behavior in the wind tunnel give confidence in the potential of this innovative idea.
Numerical optimization of circulation control airfoils
Tai, T. C.; Kidwell, G. H., Jr.; Vanderplaats, G. N.
1981-01-01
A numerical procedure for optimizing circulation control airfoils, which consists of the coupling of an optimization scheme with a viscous potential flow analysis for blowing jet, is presented. The desired airfoil is defined by a combination of three baseline shapes (cambered ellipse, and cambered ellipse with drooped and spiralled trailing edges). The coefficients of these shapes are used as design variables in the optimization process. Under the constraints of lift augmentation and lift-to-drag ratios, the optimal airfoils are found to lie between those of cambered ellipse and the drooped trailing edge, towards the latter as the angle of attack increases. Results agree qualitatively with available experimental data.
Trailing edge modifications for flatback airfoils.
Energy Technology Data Exchange (ETDEWEB)
Kahn, Daniel L. (University of California, Davis, CA); van Dam, C.P. (University of California, Davis, CA); Berg, Dale E.
2008-03-01
The adoption of blunt trailing edge airfoils (also called flatback airfoils) for the inboard region of large wind turbine blades has been proposed. Blunt trailing edge airfoils would not only provide a number of structural benefits, such as increased structural volume and ease of fabrication and handling, but they have also been found to improve the lift characteristics of thick airfoils. Therefore, the incorporation of blunt trailing edge airfoils would allow blade designers to more freely address the structural demands without having to sacrifice aerodynamic performance. These airfoils do have the disadvantage of generating high levels of drag as a result of the low-pressure steady or periodic flow in the near-wake of the blunt trailing edge. Although for rotors, the drag penalty appears secondary to the lift enhancement produced by the blunt trailing edge, high drag levels are of concern in terms of the negative effect on the torque and power generated by the rotor. Hence, devices are sought that mitigate the drag of these airfoils. This report summarizes the literature on bluff body vortex shedding and bluff body drag reduction devices and proposes four devices for further study in the wind tunnel.
Quiet airfoils for small and large wind turbines
Tangler, James L [Boulder, CO; Somers, Dan L [Port Matilda, PA
2012-06-12
Thick airfoil families with desirable aerodynamic performance with minimal airfoil induced noise. The airfoil families are suitable for a variety of wind turbine designs and are particularly well-suited for use with horizontal axis wind turbines (HAWTs) with constant or variable speed using pitch and/or stall control. In exemplary embodiments, a first family of three thick airfoils is provided for use with small wind turbines and second family of three thick airfoils is provided for use with very large machines, e.g., an airfoil defined for each of three blade radial stations or blade portions defined along the length of a blade. Each of the families is designed to provide a high maximum lift coefficient or high lift, to exhibit docile stalls, to be relatively insensitive to roughness, and to achieve a low profile drag.
Shivers, James P.
1961-01-01
Results of an investigation, conducted on the Langley helicopter test tower, of a rotor having an NACA 632A015 airfoil thickness distribution in combination with an NACA 230 mean line are presented. Comparison with a previously reported test of a symmetrical rotor blade efficiency was substantially improved over a wide range of tip Mach numbers. The maximum mean lift coefficient was essentially unchanged from that obtained with uncambered blades. Some data showing the effect of a distributed type of leading-edge roughness are also included.
Analysis for Obtaining High Lift Pressure Distributions for Transonic Airfoils.
1980-06-01
34 "Douglas Aircraft Report, MVC -J 1097/01, 1971. 3. Grabowski, W. et al., "Turbulent Flo Past a Self-PropelZed Vehicle - Part I - Formation," "Flow... Spring , MD 20910 Redstone Scientific Information Center Engineering Library Chief, Document Section McDonnell Douglas Corporation Army Missile Command
Unsteady Transonic Flow Past Airfoils in Rigid Body Motion.
1981-03-01
L( .0 0? CM - .:436 Fiqu ~r(- Sa 77 -- -- .- - --. - PiRF~I~JOUKOWSKI ARFOIL RjA- 0. 00 N- _Q C C-EfP . TI mr= 0.,00 AL[FA= 0.00 M FQ= 0.00 CEFA. 0.00...CL .526-4C C 04 CM - .22 C 2-4 Fique 1cL 108 t AL C.A I, " .. mIt oCN o:f : I mt 4 8 il .o . o .. T1 36.65 A 0.00 m ,,o 630.00 OUPA L., OOD. 000 -650C
Lawing, P. L.
1985-01-01
A method of constructing airfoils by inscribing pressure channels on the face of opposing plates, bonding them together to form one plate with integral channels, and contour machining this plate to form an airfoil model is described. The research and development program to develop the bonding technology is described as well as the construction and testing of an airfoil model. Sample aerodynamic data sets are presented and discussed. Also, work currently under way to produce thin airfoils with camber is presented. Samples of the aft section of a 6 percent airfoil with complete pressure instrumentation including the trailing edge are pictured and described. This technique is particularly useful in fabricating models for transonic cryogenic testing, but it should find application in a wide ange of model construction projects, as well as the fabrication of fuel injectors, space hardware, and other applications requiring advanced bonding technology and intricate fluid passages.
Silva, Walter A.
1993-01-01
The presentation begins with a brief description of the motivation and approach that has been taken for this research. This will be followed by a description of the Volterra Theory of Nonlinear Systems and the CAP-TSD code which is an aeroelastic, transonic CFD (Computational Fluid Dynamics) code. The application of the Volterra theory to a CFD model and, more specifically, to a CAP-TSD model of a rectangular wing with a NACA 0012 airfoil section will be presented.
Plasma Flow Control Optimized Airfoil
Voikov, Vladimir; Patel, Mehul
2005-11-01
Recent advances in flow control research have demonstrated that plasma actuators can be efficient in different aerodynamic applications, particularly in providing flight control without conventional moving surfaces. The concept involves the use of a laminar airfoil design that employs a separation ramp at the trailing edge that can be manipulated by a plasma actuator to control lift, similar to trailing-edge flaps. The advantages are lower drag by a combination of the laminar flow design, and elimination of parasitic drag associated with wing-flap junctions. This work involves numerical simulations and experiments on a HSNLF(1)-0213 airfoil. The numerical results are obtained using an unsteady, compressible Navier-Stokes simulation that includes a model for the plasma actuators. The experiments are performed on a 2-D airfoil section that is mounted on a lift-drag force balance. The results demonstrate lift enhancement produced by the plasma actuator that is comparable to a plane flap. They also reveal an optimum actuator unsteady frequency that scales with the length of the separated region and local velocity, and is associated with the generation of a train of spanwise vortices. Other scaling including the effect of Reynolds number is presented.
Illustration of airfoil shape effect on forward-swept wing divergence
Bland, S. R.
1980-01-01
A static aeroelastic analysis is presented of the divergence of untapered wings with conventional and supercritical airfoil sections at sweep angles of zero and -15 deg. One bending and one torsion mode were employed for a uniform rectangular cantilevered beam with the elastic axis at midchord, and calculations were based on a two-dimensional differential equations formulation in the structural coordinate system and in simple strip theory. A minimum divergence speed in the transonic range is obtained which is associated with the rearward shift of the aerodynamic center, and a 17% difference in minimum divergence dynamic pressure is found between a supercritical and a conventional wing. It is noted that although the strip method employed allows the assessment of the sensitivity of airfoil shapes to divergence, three-dimensional transonic aerodynamic methods should be used to predict wing divergence characteristics.
National Aeronautics and Space Administration — Techsburg is teaming with the Vibration and Acoustics Laboratory of Virginia Tech to propose a non-linear analytical tool for designing Herschel-Quincke (HQ)...
Reversible airfoils for stopped rotors in high speed flight
International Nuclear Information System (INIS)
Niemiec, Robert; Jacobellis, George; Gandhi, Farhan
2014-01-01
This study starts with the design of a reversible airfoil rib for stopped-rotor applications, where the sharp trailing-edge morphs into the rounded leading-edge, and vice-versa. A NACA0012 airfoil is approximated in a piecewise linear manner and straight, rigid outer profile links used to define the airfoil contour. The end points of the profile links connect to control links, each set on a central actuation rod via an offset. Chordwise motion of the actuation rod moves the control and the profile links and reverses the airfoil. The paper describes the design methodology and evolution of the final design, based on which two reversible airfoil ribs were fabricated and used to assemble a finite span reversible rotor/wing demonstrator. The profile links were connected by Aluminum strips running in the spanwise direction which provided stiffness as well as support for a pre-tensioned elastomeric skin. An inter-rib connector with a curved-front nose piece supports the leading-edge. The model functioned well and was able to reverse smoothly back-and-forth, on application and reversal of a voltage to the motor. Navier–Stokes CFD simulations (using the TURNS code) show that the drag coefficient of the reversible airfoil (which had a 13% maximum thickness due to the thickness of the profile links) was comparable to that of the NACA0013 airfoil. The drag of a 16% thick elliptical airfoil was, on average, about twice as large, while that of a NACA0012 in reverse flow was 4–5 times as large, even prior to stall. The maximum lift coefficient of the reversible airfoil was lower than the elliptical airfoil, but higher than the NACA0012 in reverse flow operation. (paper)
DEFF Research Database (Denmark)
Chougule, Prasad; Nielsen, Søren R.K.
2014-01-01
Nowadays, small vertical axis wind turbines are receiving more attention due to their suitability in micro-electricity generation. There are few vertical axis wind turbine designs with good power curve. However, the efficiency of power extraction has not been improved. Therefore, an attempt has...... been made to utilize high lift technology for vertical axis wind turbines in order to improve power efficiency. High lift is obtained by double-element airfoil mainly used in aeroplane wing design. In this current work a low Reynolds number airfoil is selected to design a double-element airfoil blade...... for use in vertical axis wind turbine to improve the power efficiency. Double-element airfoil blade design consists of a main airfoil and a slat airfoil. Orientation of slat airfoil is a parameter of investigation in this paper and air flow simulation over double-element airfoil. With primary wind tunnel...
Complex configuration analysis at transonic speeds
Boppe, C. W.; Aidala, P. V.
1980-01-01
Advanced performance requirements of new combat and transport aircraft together with design time constraints intensify the development and application of three dimensional computational analyses. A computational method which was developed for the specific purpose of providing an engineering analysis of complex aircraft configurations at transonic speeds. Particular attention is given to the recently incorporated wing viscous interaction and canard capabilities. The treatment of fuselage fairings, nacelles, and pylons is reviewed. The means for keeping computing resources at reasonable levels are identified. Three configurations were selected for correlations with experimental data. Taken together, the comparisons illustrate the full extent of current analysis capabilities. The configurations include: (1) a wing fuselage canard fighter; (2) a transport with fuselage fairings, four nacelles, four pylons; and (3) a space vehicle which includes an external fuel tank and rocket boosters (transonic launch configuration).
Webster, T. J.
1982-01-01
The 0.3 m Transonic Cryogenic Tunnel (TCT) at the NASA Langley Research Center was built in 1973 as a facility intended to be used for no more than 60 hours in order to verify the validity of the cryogenic wind tunnel concept at transonic speeds. The role of the 0.3 m TCT has gradually changed until now, after over 3000 hours of operation, it is classified as a major NASA research facility and, under the administration of the Experimental Techniques Branch, it is used extensively for the testing of airfoils at high Reynolds numbers and for the development of various technologies related to the efficient operation and use of cryogenic wind tunnels. The purpose of this report is to document the results of a recent safety analysis of the 0.3 m TCT facility. This analysis was made as part of an on going program with the Experimental Techniques Branch designed to ensure that the existing equipment and current operating procedures of the 0.3 m TCT facility are acceptable in terms of today's standards of safety for cryogenic systems.
Schnell, W. C.; Ordonez, G. W.
1981-01-01
A 1/8 scale jet-effects model was tested in the NASA Ames 11 ft transonic tunnel at static conditions and over a range of Mach numbers from 0.4 to 1.4. The data presented show that significant differences in aeropropulsion performance can be expected by varying the exhaust nozzle type and its geometric parameters on a V/STOL underwing nacelle installation.
Morren, Sybil Huang
1991-01-01
Transonic flow of dense gases for two-dimensional, steady-state, flow over a NACA 0012 airfoil was predicted analytically. The computer code used to model the dense gas behavior was a modified version of Jameson's FL052 airfoil code. The modifications to the code enabled modeling the dense gas behavior near the saturated vapor curve and critical pressure region where the fundamental derivative, Gamma, is negative. This negative Gamma region is of interest because the nonclassical gas behavior such as formation and propagation of expansion shocks, and the disintegration of inadmissible compression shocks may exist. The results indicated that dense gases with undisturbed thermodynamic states in the negative Gamma region show a significant reduction in the extent of the transonic regime as compared to that predicted by the perfect gas theory. The results support existing theories and predictions of the nonclassical, dense gas behavior from previous investigations.
Wind turbine airfoil catalogue
DEFF Research Database (Denmark)
Bertagnolio, F.; Sørensen, Niels N.; Johansen, Jeppe
2001-01-01
The aim of this work is two-sided. Firstly, experimental results obtained for numerous sets of airfoil measurements (mainly intended for wind turbine applications) are collected and compared with computational results from the 2D Navier-Stokes solverEllipSys2D, as well as results from the panel...... data. A study correlating the available data and this classification is performed. It is found that transition modelling is to a large extent responsible forthe poor quality of the computational results for most of the considered airfoils. The transition model mechanism that leads...
Wind turbine airfoil catalogue
Bertagnolio, F.; Sørensen, Niels N.; Johansen, Jeppe; Fuglsang, P.
2001-01-01
The aim of this work is two-sided. Firstly, experimental results obtained for numerous sets of airfoil measurements (mainly intended for wind turbine applications) are collected and compared with computational results from the 2D Navier-Stokes solverEllipSys2D, as well as results from the panel method code XFOIL. Secondly, we are interested in validating the code EllipSys2D and finding out for which airfoils it does not perform well compared to the experiments, as well as why, when it does so...
High-efficiency airfoil rudders applied to submarines
Directory of Open Access Journals (Sweden)
ZHOU Yimei
2017-03-01
Full Text Available Modern submarine design puts forward higher and higher requirements for control surfaces, and this creates a requirement for designers to constantly innovate new types of rudder so as to improve the efficiency of control surfaces. Adopting the high-efficiency airfoil rudder is one of the most effective measures for improving the efficiency of control surfaces. In this paper, we put forward an optimization method for a high-efficiency airfoil rudder on the basis of a comparative analysis of the various strengths and weaknesses of the airfoil, and the numerical calculation method is adopted to analyze the influence rule of the hydrodynamic characteristics and wake field by using the high-efficiency airfoil rudder and the conventional NACA rudder comparatively; at the same time, a model load test in a towing tank was carried out, and the test results and simulation calculation obtained good consistency:the error between them was less than 10%. The experimental results show that the steerage of a high-efficiency airfoil rudder is increased by more than 40% when compared with the conventional rudder, but the total resistance is close:the error is no more than 4%. Adopting a high-efficiency airfoil rudder brings much greater lifting efficiency than the total resistance of the boat. The results show that high-efficiency airfoil rudder has obvious advantages for improving the efficiency of control, giving it good application prospects.
Agopian, K. G.
1974-01-01
The problem of inviscid, steady transonic conical flow, formulated in terms of the small disturbance theory, is studied. The small disturbance equation and similarity rules are presented, and a boundary value problem is formulated for the case of a supersonic freestream Mach number. The equation for the perturbation potential is solved numerically using an elliptic finite difference system. The difference equations are solved with a point relaxation algorithm that is also capable of capturing the shock wave during the iteration procedure by using the boundary conditions at the shock. Numerical calculations, for shock location, pressure distribution and drag coefficient, are presented for a family of nonlifting conical wings. The theory of slender wings is also presented and analytical results for pressure and drag coefficients are obtained.
Numerical simulation of the RISOe1-airfoil dynamic stall
Energy Technology Data Exchange (ETDEWEB)
Bertagnolio, F.; Soerensen, N. [Risoe National Lab., Wind Energy and Atmospheric Physics Dept., Roskilde (Denmark)
1997-12-31
In this paper we are concerned with the numerical computation of the dynamic stall that occur in the viscous flowfield over an airfoil. These results are compared to experimental data that were obtained with the new designed RISOe1-airfoil, both for a motionless airfoil and for a pitching motion. Moreover, we present some numerical computations of the plunging and lead-lag motions. We also investigate the possibility of using the pitching motion to simulate the plunging and lead-lag situations. (au)
S833, S834, and S835 Airfoils: November 2001--November 2002
Energy Technology Data Exchange (ETDEWEB)
Somers, D. M.
2005-08-01
A family of quiet, thick, natural-laminar-flow airfoils, the S833, S834, and S835, for 1 - 3-meter-diameter, variable-speed/variable-pitch, horizontal-axis wind turbines has been designed and analyzed theoretically. The two primary objectives of high maximum lift, relatively insensitive to roughness, and low profile drag have been achieved. The airfoils should exhibit docile stalls, which meet the design goal. The constraints on the pitching moment and the airfoils thicknesses have been satisfied.
S830, S831, and S832 Airfoils: November 2001-November 2002
Energy Technology Data Exchange (ETDEWEB)
Somers, D. M.
2005-08-01
A family of quiet, thick, natural-laminar-flow airfoils, the S830, S831, and S832, for 40 - 50-meter-diameter, variable-speed/variable-pitch, horizontal-axis wind turbines has been designed and analyzed theoretically. The two primary objectives of high maximum lift, relatively insensitive to roughness, and low profile drag have been achieved. The airfoils should exhibit docile stalls, which meet the design goal. The constraints on the pitching moment and the airfoils thicknesses have been satisfied.
DEFF Research Database (Denmark)
Chougule, Prasad; Rosendahl, Lasse; Nielsen, Søren R.K.
2015-01-01
A design of double-element airfoil is proposed for its use in the vertical axis wind turbine. The double-element airfoil system consists of a main airfoil and a slat airfoil. The design parameters of the double-element airfoil system are given by the position and orientation of the trailing edge...... of the slat airfoil with respect to the nose of the main airfoil. The position parameters are given by xTE and yTE. The orientation is given by the slat angle βs; it is the angle between the chord lengths of the slat airfoil with respect to the main airfoil. Various tests were performed using different slat...
Steckel, D. K.; Dahlin, J. A.; Henne, P. A.
1980-01-01
These basic characteristics of critical wings included wing area, aspect ratio, average thickness, and sweep as well as practical constraints on the planform and thickness near the wing root to allow for the landing gear. Within these constraints, a large matrix of wing designs was studied with spanwise variations in the types of airfoils and distribution of lift as well as some small planform changes. The criteria by which the five candidate wings were chosen for testing were the cruise and buffet characteristics in the transonic regime and the compatibility of the design with low speed (high-lift) requirements. Five wing-wide-body configurations were tested in the NASA Ames 11-foot transonic wind tunnel. Nacelles and pylons, flap support fairings, tail surfaces, and an outboard aileron were also tested on selected configurations.
Vertical axis wind turbine airfoil
Krivcov, Vladimir; Krivospitski, Vladimir; Maksimov, Vasili; Halstead, Richard; Grahov, Jurij Vasiljevich
2012-12-18
A vertical axis wind turbine airfoil is described. The wind turbine airfoil can include a leading edge, a trailing edge, an upper curved surface, a lower curved surface, and a centerline running between the upper surface and the lower surface and from the leading edge to the trailing edge. The airfoil can be configured so that the distance between the centerline and the upper surface is the same as the distance between the centerline and the lower surface at all points along the length of the airfoil. A plurality of such airfoils can be included in a vertical axis wind turbine. These airfoils can be vertically disposed and can rotate about a vertical axis.
Periodic transonic flow simulation using fourier-based algorithm
International Nuclear Information System (INIS)
Mohaghegh, Mohammad Reza; Malekjafarian, Majid
2014-01-01
The present research simulates time-periodic unsteady transonic flow around pitching airfoils via the solution of unsteady Euler and Navier-Stokes equations, using time spectral method (TSM) and compares it with the traditional methods like BDF and explicit structured adaptive grid method. The TSM uses a Fourier representation in time and hence solves for the periodic state directly without resolving transients (which consume most of the resources in a time-accurate scheme). Mathematical tools used here are discrete Fourier transformations. The TSM has been validated with 2D external aerodynamics test cases. These test cases are NACA 64A010 (CT6) and NACA 0012 (CT1 and CT5) pitching airfoils. Because of turbulent nature of flow, Baldwin-Lomax turbulence model has been used in viscous flow analysis with large oscillation amplitude (CT5 type). The results presented by the TSM are compared with experimental data and the two other methods. By enforcing periodicity and using Fourier representation in time that has a spectral accuracy, tremendous reduction of computational cost has been obtained compared to the conventional time-accurate methods. Results verify the small number of time intervals per pitching cycle (just four time intervals) required to capture the flow physics with small oscillation amplitude (CT6) and large oscillation amplitude (CT5) as compared to the other two methods.
1976-02-18
thickness fin mounted on a cylindrical body reduces to Spreiter’s earlier works [36, 37] and to that of Lawrence and Flax [L~2] • 10 Figure 22...supposed that: = 1 + f(M ) ret s s (69) where the coefficient, r, is determined empirically, and relates to the N factor of Reference 61. The function f...Transonic Flows About Certain Classes of Airfoils and Slender Bodies, NASA CR-1722, April 1971. 42. Lawrence, H. R., and Flax , A. H., "Body-Wing
Application of the Green's function method for 2- and 3-dimensional steady transonic flows
Tseng, K.
1984-01-01
A Time-Domain Green's function method for the nonlinear time-dependent three-dimensional aerodynamic potential equation is presented. The Green's theorem is being used to transform the partial differential equation into an integro-differential-delay equation. Finite-element and finite-difference methods are employed for the spatial and time discretizations to approximate the integral equation by a system of differential-delay equations. Solution may be obtained by solving for this nonlinear simultaneous system of equations in time. This paper discusses the application of the method to the Transonic Small Disturbance Equation and numerical results for lifting and nonlifting airfoils and wings in steady flows are presented.
2011-05-13
sound production from a hydrofoil and identified three mechanisms: (1) low frequency curvature noise associated with interaction of a turbulent...2002). 2 Technical Approach A two-dimensional, dual-slotted, elliptic circulation control airfoil based on the hydrofoil studied by Rogers...airfoil, shown in Figure 1A, is designed based on the geometry of the hydrofoil previously studied by Rogers & Donnelly (2004). The airfoil’s profile
Tail Rotor Airfoils Stabilize Helicopters, Reduce Noise
2010-01-01
Founded by former Ames Research Center engineer Jim Van Horn, Van Horn Aviation of Tempe, Arizona, built upon a Langley Research Center airfoil design to create a high performance aftermarket tail rotor for the popular Bell 206 helicopter. The highly durable rotor has a lifetime twice that of the original equipment manufacturer blade, reduces noise by 40 percent, and displays enhanced performance at high altitudes. These improvements benefit helicopter performance for law enforcement, military training, wildfire and pipeline patrols, and emergency medical services.
Airfoil characteristics for wind turbines
DEFF Research Database (Denmark)
Bak, C.; Fuglsang, P.; Sørensen, Niels N.
1999-01-01
is capable of determining the correct qualitative behaviour for airfoils subject to rotation. The method shows that lift is high at the root compared to 2D airfoil characteristics. Thedifferent systematic methods show the importance of rotational and 3D effects on rotors. Furthermore, the methods show high......Airfoil characteristics for use in the Blade Element Momentum (BEM) method calculating the forces on Horizontal Axis Wind Turbines (HAWT) are derived by use of systematic methods. The investigation and derivation of the airfoil characteristics are basedon four different methods: 1) Inverse momentum...... theory, 2) Actuator disc theory, 3) Numerical optimisation and 4) Quasi-3D CFD computations. The two former methods are based on 3D CFD computations and wind tunnel measurements on a 41-m full-scale rotorwith LM 19.1 blades. The derived airfoil characteristics show that the lift coefficient in stall...
Flatback airfoil wind tunnel experiment.
Energy Technology Data Exchange (ETDEWEB)
Mayda, Edward A. (University of California, Davis, CA); van Dam, C.P. (University of California, Davis, CA); Chao, David D. (University of California, Davis, CA); Berg, Dale E.
2008-04-01
A computational fluid dynamics study of thick wind turbine section shapes in the test section of the UC Davis wind tunnel at a chord Reynolds number of one million is presented. The goals of this study are to validate standard wind tunnel wall corrections for high solid blockage conditions and to reaffirm the favorable effect of a blunt trailing edge or flatback on the performance characteristics of a representative thick airfoil shape prior to building the wind tunnel models and conducting the experiment. The numerical simulations prove the standard wind tunnel corrections to be largely valid for the proposed test of 40% maximum thickness to chord ratio airfoils at a solid blockage ratio of 10%. Comparison of the computed lift characteristics of a sharp trailing edge baseline airfoil and derived flatback airfoils reaffirms the earlier observed trend of reduced sensitivity to surface contamination with increasing trailing edge thickness.
Leading-edge slat optimization for maximum airfoil lift
Olson, L. E.; Mcgowan, P. R.; Guest, C. J.
1979-01-01
A numerical procedure for determining the position (horizontal location, vertical location, and deflection) of a leading edge slat that maximizes the lift of multielement airfoils is presented. The structure of the flow field is calculated by iteratively coupling potential flow and boundary layer analysis. This aerodynamic calculation is combined with a constrained function minimization analysis to determine the position of a leading edge slat so that the suction peak on the nose of the main airfoil is minized. The slat position is constrained by the numerical procedure to ensure an attached boundary layer on the upper surface of the slat and to ensure negligible interaction between the slat wake and the boundary layer on the upper surface of the main airfoil. The highest angle attack at which this optimized slat position can maintain attached flow on the main airfoil defines the optimum slat position for maximum lift. The design method is demonstrated for an airfoil equipped with a leading-edge slat and a trailing edge, single-slotted flap. The theoretical results are compared with experimental data, obtained in the Ames 40 by 80 Foot Wind Tunnel, to verify experimentally the predicted slat position for maximum lift. The experimentally optimized slat position is in good agreement with the theoretical prediction, indicating that the theoretical procedure is a feasible design method.
Airfoil Shape Optimization based on Surrogate Model
Mukesh, R.; Lingadurai, K.; Selvakumar, U.
2018-02-01
Engineering design problems always require enormous amount of real-time experiments and computational simulations in order to assess and ensure the design objectives of the problems subject to various constraints. In most of the cases, the computational resources and time required per simulation are large. In certain cases like sensitivity analysis, design optimisation etc where thousands and millions of simulations have to be carried out, it leads to have a life time of difficulty for designers. Nowadays approximation models, otherwise called as surrogate models (SM), are more widely employed in order to reduce the requirement of computational resources and time in analysing various engineering systems. Various approaches such as Kriging, neural networks, polynomials, Gaussian processes etc are used to construct the approximation models. The primary intention of this work is to employ the k-fold cross validation approach to study and evaluate the influence of various theoretical variogram models on the accuracy of the surrogate model construction. Ordinary Kriging and design of experiments (DOE) approaches are used to construct the SMs by approximating panel and viscous solution algorithms which are primarily used to solve the flow around airfoils and aircraft wings. The method of coupling the SMs with a suitable optimisation scheme to carryout an aerodynamic design optimisation process for airfoil shapes is also discussed.
Turbine Airfoil Leading Edge Film Cooling Bibliography: 1972–1998
Directory of Open Access Journals (Sweden)
D. M. Kercher
2000-01-01
Full Text Available Film cooling for turbine airfoil leading edges has been a common practice for at least 35 years as turbine inlet gas temperatures and pressures have continually increased along with cooling air temperatures for higher engine cycle efficiency. With substantial engine cycle performance improvements from higher gas temperatures, it has become increasingly necessary to film cool nozzle and rotor blade leading edges since external heat transfer coefficients and thus heat load are the highest in this airfoil region. Optimum cooling air requirements in this harsh environment has prompted a significant number of film cooling investigations and analytical studies reported over the past 25 years from academia, industry and government agencies. Substantial progress has been made in understanding the complex nature of leading edge film cooling from airfoil cascades, simulated airfoil leading edges and environment. This bibliography is a report of the open-literature references available which provide information on the complex aero–thermo interaction of leading edge gaseous film cooling with mainstream flow. From much of this investigative information has come successful operational leading edge film cooling design systems capable of sustaining airfoil leading edge durability in very hostile turbine environments.
Tables of properties of airfoil polynomials
Desmarais, Robert N.; Bland, Samuel R.
1995-01-01
This monograph provides an extensive list of formulas for airfoil polynomials. These polynomials provide convenient expansion functions for the description of the downwash and pressure distributions of linear theory for airfoils in both steady and unsteady subsonic flow.
Pressure distribution over NACA 23012 airfoil with a slotted and a split flap
Harris, Thomas A; Lowry, John G
1941-01-01
A pressure-distribution investigation has been conducted in the NACA 4 by 6-foot vertical wind tunnel to determine the air loads on an NACA 23012 airfoil in combination with a 25.66-percent-chord slotted flap and a 20-percent-chord split flap. Pressures were measured on both the upper and the lower surfaces of the main airfoil and the flaps for several angles of attack and at several flap settings. The data, presented as pressure diagrams and as graphs of the section coefficients for the flap alone and for the airfoil-flap combinations, are applicable to rib and flap design for a combination of a thick airfoil and a slotted or a split flap. The results of previous tests of a NACA 23012 airfoil with a slotted flap are compared with the present results.
Characterization of Oscillatory Lift in MFC Airfoils
Lang Jr, Joseph Reagle
2014-01-01
The purpose of this research is to characterize the response of an airfoil with an oscillatory morphing, Macro-fiber composite (MFC) trailing edge. Correlation of the airfoil lift with the oscillatory input is presented. Modal analysis of the test airfoil and apparatus is used to determine the frequency response function. The effects of static MFC inputs on the FRF are presented and compared to the unactuated airfoil. The transfer function is then used to determine the lift component du...
Linear Strength Vortex Panel Method for NACA 4412 Airfoil
Liu, Han
2018-03-01
The objective of this article is to formulate numerical models for two-dimensional potential flow over the NACA 4412 Airfoil using linear vortex panel methods. By satisfying the no penetration boundary condition and Kutta condition, the circulation density on each boundary points (end point of every panel) are obtained and according to which, surface pressure distribution and lift coefficients of the airfoil are predicted and validated by Xfoil, an interactive program for the design and analysis of airfoil. The sensitivity of results to the number of panels is also investigated in the end, which shows that the results are sensitive to the number of panels when panel number ranges from 10 to 160. With the increasing panel number (N>160), the results become relatively insensitive to it.
Trailing Edge Noise Model Validation and Application to Airfoil Optimization
DEFF Research Database (Denmark)
Bertagnolio, Franck; Aagaard Madsen, Helge; Bak, Christian
2010-01-01
The aim of this article is twofold. First, an existing trailing edge noise model is validated by comparing with airfoil surface pressure fluctuations and far field sound pressure levels measured in three different experiments. The agreement is satisfactory in one case but poor in two other cases...... noise emission, trying at the same time to preserve some of its aerodynamic and geometric characteristics. The new designs are characterized by less cambered airfoils and flatter suction sides. The resulting noise reductions seem to be mainly achieved by a reduction in the turbulent kinetic energy...
Root region airfoil for wind turbine
Tangler, James L.; Somers, Dan M.
1995-01-01
A thick airfoil for the root region of the blade of a wind turbine. The airfoil has a thickness in a range from 24%-26% and a Reynolds number in a range from 1,000,000 to 1,800,000. The airfoil has a maximum lift coefficient of 1.4-1.6 that has minimum sensitivity to roughness effects.
Internal flow measurement in transonic compressor by PIV technique
Wang, Tongqing; Wu, Huaiyu; Liu, Yin
2001-11-01
The paper presents some research works conducted in National Key Laboratory of Aircraft Engine of China on the shock containing supersonic flow measurement as well as the internal flow measurement of transoijc compressor by PIC technique. A kind of oil particles in diameter about 0.3 micrometers containing in the flow was discovered to be a very good seed for the PIV measurement of supersonic jet flow. The PIV measurement in over-expanded supersonic free jet and in the flow over wages show a very clear shock wave structure. In the PIV internal flow measurement of transonic compressor a kind of liquid particle of glycol was successful to be used as the seed. An illumination periscope with sheet forming optics was designed and manufactured, it leaded the laser shot generated from an integrate dual- cavity Nd:YAG laser of TSI PIV results of internal flow of an advanced low aspect ratio transonic compressor were shown and discussed briefly.
Directory of Open Access Journals (Sweden)
Phil Ligrani
2012-01-01
Full Text Available The influences of a variety of different physical phenomena are described as they affect the aerodynamic performance of turbine airfoils in compressible, high-speed flows with either subsonic or transonic Mach number distributions. The presented experimental and numerically predicted results are from a series of investigations which have taken place over the past 32 years. Considered are (i symmetric airfoils with no film cooling, (ii symmetric airfoils with film cooling, (iii cambered vanes with no film cooling, and (iv cambered vanes with film cooling. When no film cooling is employed on the symmetric airfoils and cambered vanes, experimentally measured and numerically predicted variations of freestream turbulence intensity, surface roughness, exit Mach number, and airfoil camber are considered as they influence local and integrated total pressure losses, deficits of local kinetic energy, Mach number deficits, area-averaged loss coefficients, mass-averaged total pressure loss coefficients, omega loss coefficients, second law loss parameters, and distributions of integrated aerodynamic loss. Similar quantities are measured, and similar parameters are considered when film-cooling is employed on airfoil suction surfaces, along with film cooling density ratio, blowing ratio, Mach number ratio, hole orientation, hole shape, and number of rows of holes.
Airfoil selection methodology for Small Wind Turbines
DEFF Research Database (Denmark)
Salgado Fuentes, Valentin; Troya, Cesar; Moreno, Gustavo
2016-01-01
a new methodology for airfoil selection used in feasibility and optimization of small wind turbines with low cut-in speed. On the first stage, airfoils data is tested on XFOIL software to check its compatibility with the simulator; then, arithmetic mean criteria is recursively used to discard...... underperformed airfoils; the best airfoil data was exported to Matlab for a deeper analysis. In the second part, data points were interpolated using "splines" to calculate glide ratio and stability across multiple angles of attack, those who present a bigger steadiness were conserved. As a result, 3 airfoils...
Airfoil shape optimization using non-traditional optimization technique and its validation
Directory of Open Access Journals (Sweden)
R. Mukesh
2014-07-01
Full Text Available Computational fluid dynamics (CFD is one of the computer-based solution methods which is more widely employed in aerospace engineering. The computational power and time required to carry out the analysis increase as the fidelity of the analysis increases. Aerodynamic shape optimization has become a vital part of aircraft design in the recent years. Generally if we want to optimize an airfoil we have to describe the airfoil and for that, we need to have at least hundred points of x and y co-ordinates. It is really difficult to optimize airfoils with this large number of co-ordinates. Nowadays many different schemes of parameter sets are used to describe general airfoil such as B-spline, and PARSEC. The main goal of these parameterization schemes is to reduce the number of needed parameters as few as possible while controlling the important aerodynamic features effectively. Here the work has been done on the PARSEC geometry representation method. The objective of this work is to introduce the knowledge of describing general airfoil using twelve parameters by representing its shape as a polynomial function. And also we have introduced the concept of Genetic Algorithm to optimize the aerodynamic characteristics of a general airfoil for specific conditions. A MATLAB program has been developed to implement PARSEC, Panel Technique, and Genetic Algorithm. This program has been tested for a standard NACA 2411 airfoil and optimized to improve its coefficient of lift. Pressure distribution and co-efficient of lift for airfoil geometries have been calculated using the Panel method. The optimized airfoil has improved co-efficient of lift compared to the original one. The optimized airfoil is validated using wind tunnel data.
A Two Element Laminar Flow Airfoil Optimized for Cruise. M.S. Thesis
Steen, Gregory Glen
1994-01-01
Numerical and experimental results are presented for a new two-element, fixed-geometry natural laminar flow airfoil optimized for cruise Reynolds numbers on the order of three million. The airfoil design consists of a primary element and an independent secondary element with a primary to secondary chord ratio of three to one. The airfoil was designed to improve the cruise lift-to-drag ratio while maintaining an appropriate landing capability when compared to conventional airfoils. The airfoil was numerically developed utilizing the NASA Langley Multi-Component Airfoil Analysis computer code running on a personal computer. Numerical results show a nearly 11.75 percent decrease in overall wing drag with no increase in stall speed at sailplane cruise conditions when compared to a wing based on an efficient single element airfoil. Section surface pressure, wake survey, transition location, and flow visualization results were obtained in the Texas A&M University Low Speed Wind Tunnel. Comparisons between the numerical and experimental data, the effects of the relative position and angle of the two elements, and Reynolds number variations from 8 x 10(exp 5) to 3 x 10(exp 6) for the optimum geometry case are presented.
Directory of Open Access Journals (Sweden)
Kan Kan
2016-01-01
Full Text Available Airfoil is a key factor that influences the hydrodynamic performance of vertical axis tidal current energy turbine. In order to explore the influences from the runner’s blade airfoil towards its hydrodynamic performance, three-dimensional unsteady flow numerical simulation on four airfoils of straight blade H type of tidal current energy turbines was carried out, under the precondition of same turbine compactness. Through investigating the influences from the four different airfoils of H blade towards the runner in terms of its dynamic torque feature and hydropower utilization coefficient, this research has analyzed the hydrodynamic performance of the tidal current energy turbine. As the research result indicates, the maximum dynamic torque value of the single-blade turbine with NACA0015 airfoil is significantly higher than that of the other three airfoils; while the maximum value of the overall runner torque is determined by the operating conditions such as different stream speed etc. As a whole, hydropower utilization coefficient of the turbine with asymmetric airfoil NACA4415 is better than that of the other three airfoils. This research has provided references for the design and airfoil selection for the new type of tidal current energy turbine.
VTOL to Transonic Aircraft, Phase I
National Aeronautics and Space Administration — The cyclogyro, an aircraft propulsion concept with the potential for VTOL to the lower bounds of transonic flight, is conceptually simple but structurally and...
An Experimental Comparison Between Flexible and Rigid Airfoils at Low Reynolds Numbers
Uzodinma, Jaylon; Macphee, David
2017-11-01
This study uses experimental and computational research methods to compare the aerodynamic performance of rigid and flexible airfoils at a low Reynolds number throughout varying angles of attack. This research can be used to improve the design of small wind turbines, micro-aerial vehicles, and any other devices that operate at low Reynolds numbers. Experimental testing was conducted in the University of Alabama's low-speed wind tunnel, and computational testing was conducted using the open-source CFD code OpenFOAM. For experimental testing, polyurethane-based (rigid) airfoils and silicone-based (flexible) airfoils were constructed using acrylic molds for NACA 0012 and NACA 2412 airfoil profiles. Computer models of the previously-specified airfoils were also created for a computational analysis. Both experimental and computational data were analyzed to examine the critical angles of attack, the lift and drag coefficients, and the occurrence of laminar boundary separation for each airfoil. Moreover, the computational simulations were used to examine the resulting flow fields, in order to provide possible explanations for the aerodynamic performances of each airfoil type. EEC 1659710.
Yaros, Steven F.; Carlson, John R.; Chandrasekaran, Balasubramanyan
1986-01-01
An effort has been undertaken at the NASA Langley Research Center to assess the capabilities of available computational methods for use in propulsion integration design studies of transonic transport aircraft, particularly of pylon/nacelle combinations which exhibit essentially no interference drag. The three computer codes selected represent state-of-the-art computational methods for analyzing complex configurations at subsonic and transonic flight conditions. These are: EULER, a finite volume solution of the Euler equation; VSAERO, a panel solution of the Laplace equation; and PPW, a finite difference solution of the small disturbance transonic equations. In general, all three codes have certain capabilities that allow them to be of some value in predicting the flows about transport configurations, but all have limitations. Until more accurate methods are available, careful application and interpretation of the results of these codes are needed.
Yaros, S. F.; Carlson, J. R.; Chandrasekaran, B.
1986-01-01
An effort has been undertaken at the NASA Langley Research Center to assess the capabilities of available computational methods for use in propulsion integration design studies of transonic transport aircraft, particularly of pylon/nacelle combinations which exhibit essentially no interference drag. The three computer codes selected represent state-of-the-art computational methods for analyzing complex configurations at subsonic and transonic flight conditions. These are: EULER, a finitie volume solution of the Euler equation; VSAERO, a panel solution of the Laplace equation; and PPW, a finite difference solution of the small disturbance transonic equations. In general, all three codes have certain capabilities that allow them to be of some value in predicting the flows about transport configurations, but all have limitations. Until more accurate methods are available, careful application and interpretation of the results of these codes are needed.
Flow Control in a Transonic Diffuser
Gartner, Jeremy; Amitay, Michael
2014-11-01
In some airplanes such as fighter jets and UAV, short inlet ducts replace the more conventional ducts due to their shorter length. However, these ducts are associated with low length-to-diameter ratio and low aspect ratio and, thus, experience massive separation and the presence of secondary flow structures. These flow phenomena are undesirable as they lead to pressure losses and distortion at the Aerodynamic Interface Plane (AIP), where the engine face is located. It causes the engine to perform with a lower efficiency as it would with a straight duct diffuser. Different flow control techniques were studied on the short inlet duct, with the goal to reattach the flow and minimize the distortions at the AIP. Due to the complex interaction between the separation and the secondary flow structures, the necessity to understand the flow mechanisms, and how to control them at a more fundamental level, a new transonic diffuser with an upper ramp and a straight floor was designed and built. The objective of this project is to explore the effectiveness of different flow control techniques in a high subsonic (up to Mach 0.8) diffuser, so that the quasi two-dimensional separation and the formation of secondary flow structure can be isolated using a canonical flow field. Supported by Northrop Grumman.
Monitoring pressure profiles across an airfoil with a fiber Bragg grating sensor array
Papageorgiou, Anthony W.; Parkinson, Luke A.; Karas, Andrew R.; Hansen, Kristy L.; Arkwright, John W.
2018-02-01
Fluid flow over an airfoil section creates a pressure difference across the upper and lower surfaces, thus generating lift. Successful wing design is a combination of engineering design and experience in the field, with subtleties in design and manufacture having significant impact on the amount of lift produced. Current methods of airfoil optimization and validation typically involve computational fluid dynamics (CFD) and extensive wind tunnel testing with pressure sensors embedded into the airfoil to measure the pressure over the wing. Monitoring pressure along an airfoil in a wind tunnel is typically achieved using surface pressure taps that consist of hollow tubes running from the surface of the airfoil to individual pressure sensors external to the tunnel. These pressure taps are complex to configure and not ideal for in-flight testing. Fiber Bragg grating (FBG) pressure sensing arrays provide a highly viable option for both wind tunnel and inflight pressure measurement. We present a fiber optic sensor array that can detect positive and negative pressure suitable for validating CFD models of airfoil profile sections. The sensing array presented here consists of 6 independent sensing elements, each capable of a pressure resolution of less than 10 Pa over the range of 70 kPa to 120 kPa. The device has been tested with the sensor array attached to a 90mm chord length airfoil section subjected to low velocity flow. Results show that the arrays are capable of accurately detecting variations of the pressure profile along the airfoil as the angle of attack is varied from zero to the point at which stall occurs.
Numerical Simulations of Subscale Wind Turbine Rotor Inboard Airfoils at Low Reynolds Number
Energy Technology Data Exchange (ETDEWEB)
Blaylock, Myra L. [Sandia National Lab. (SNL-CA), Livermore, CA (United States). Thermal/ Fluid Sciences & Engineering Dept.; Maniaci, David Charles [Sandia National Lab. (SNL-NM), Albuquerque, NM (United States). Wind Energy Technologies Dept.; Resor, Brian R. [Sandia National Lab. (SNL-NM), Albuquerque, NM (United States). Wind Energy Technologies Dept.
2015-04-01
New blade designs are planned to support future research campaigns at the SWiFT facility in Lubbock, Texas. The sub-scale blades will reproduce specific aerodynamic characteristics of utility-scale rotors. Reynolds numbers for megawatt-, utility-scale rotors are generally above 2-8 million. The thickness of inboard airfoils for these large rotors are typically as high as 35-40%. The thickness and the proximity to three-dimensional flow of these airfoils present design and analysis challenges, even at the full scale. However, more than a decade of experience with the airfoils in numerical simulation, in the wind tunnel, and in the field has generated confidence in their performance. Reynolds number regimes for the sub-scale rotor are significantly lower for the inboard blade, ranging from 0.7 to 1 million. Performance of the thick airfoils in this regime is uncertain because of the lack of wind tunnel data and the inherent challenge associated with numerical simulations. This report documents efforts to determine the most capable analysis tools to support these simulations in an effort to improve understanding of the aerodynamic properties of thick airfoils in this Reynolds number regime. Numerical results from various codes of four airfoils are verified against previously published wind tunnel results where data at those Reynolds numbers are available. Results are then computed for other Reynolds numbers of interest.
An airloads theory for morphing airfoils in dynamic stall with experimental correlation
Ahaus, Loren A.
Helicopter rotor blades frequently encounter dynamic stall during normal flight conditions, limiting the applicability of classical thin-airfoil theory at large angles of attack. Also, it is evident that because of the largely different conditions on the advancing and retreating sides of the rotor, future rotorcraft may incorporate dynamically morphing airfoils (trailing-edge aps, dynamic camber, dynamic droop, etc.). Reduced-order aerodynamic models are needed for preliminary design and ight simulation. A unified model for predicting the airloads on a morphing airfoil in dynamic stall is presented, consisting of three components. First, a linear airloads theory allows for arbitrary airfoil deformations consistent with a morphing airfoil. Second, to capture the effects of the wake, the airloads theory is coupled to an induced ow model. Third, the overshoot and time delay associated with dynamic stall are modeled by a second-order dynamic filter, along the lines of the ONERA dynamic stall model. This paper presents a unified airloads model that allows arbitrary airfoil morphing with dynamic stall. Correlations with experimental data validate the theory.
Leading edge embedded fan airfoil concept -- A new powered high lift technology
Phan, Nhan Huu
A new powered-lift airfoil concept called Leading Edge Embedded Fan (LEEF) is proposed for Extremely Short Take-Off and Landing (ESTOL) and Vertical Take-Off and Landing (VTOL) applications. The LEEF airfoil concept is a powered-lift airfoil concept capable of generating thrust and very high lift-coefficient at extreme angles-of attack (AoA). It is designed to activate only at the take-off and landing phases, similar to conventional flaps or slats, allowing the aircraft to operate efficiently at cruise in its conventional configuration. The LEEF concept consists of placing a crossflow fan (CFF) along the leading-edge (LE) of the wing, and the housing is designed to alter the airfoil shape between take-off/landing and cruise configurations with ease. The unique rectangular cross section of the crossflow fan allows for its ease of integration into a conventional subsonic wing. This technology is developed for ESTOL aircraft applications and is most effectively applied to General Aviation (GA) aircraft. Another potential area of application for LEEF is tiltrotor aircraft. Unlike existing powered high-lift systems, the LEEF airfoil uses a local high-pressure air source from cross-flow fans, does not require ducting, and is able to be deployed using distributed electric power systems throughout the wing. In addition to distributed lift augmentation, the LEEF system can provide additional thrust during takeoff and landing operation to supplement the primary cruise propulsion system. Two-dimensional (2D) and three-dimensional (3D) Computational Fluid Dynamics (CFD) simulations of a conventional airfoil/wing using the NACA 63-3-418 section, commonly used in GA, and a LEEF airfoil/wing embedded into the same airfoil section were carried out to evaluate the advantages of and the costs associated with implementing the LEEF concept. Computational results show that significant lift and augmented thrust are available during LEEF operation while requiring only moderate fan power
Theory and Low-Order Modeling of Unsteady Airfoil Flows
Ramesh, Kiran
Unsteady flow phenomena are prevalent in a wide range of problems in nature and engineering. These include, but are not limited to, aerodynamics of insect flight, dynamic stall in rotorcraft and wind turbines, leading-edge vortices in delta wings, micro-air vehicle (MAV) design, gust handling and flow control. The most significant characteristics of unsteady flows are rapid changes in the circulation of the airfoil, apparent-mass effects, flow separation and the leading-edge vortex (LEV) phenomenon. Although experimental techniques and computational fluid dynamics (CFD) methods have enabled the detailed study of unsteady flows and their underlying features, a reliable and inexpensive loworder method for fast prediction and for use in control and design is still required. In this research, a low-order methodology based on physical principles rather than empirical fitting is proposed. The objective of such an approach is to enable insights into unsteady phenomena while developing approaches to model them. The basis of the low-order model developed here is unsteady thin-airfoil theory. A time-stepping approach is used to solve for the vorticity on an airfoil camberline, allowing for large amplitudes and nonplanar wakes. On comparing lift coefficients from this method against data from CFD and experiments for some unsteady test cases, it is seen that the method predicts well so long as LEV formation does not occur and flow over the airfoil is attached. The formation of leading-edge vortices (LEVs) in unsteady flows is initiated by flow separation and the formation of a shear layer at the airfoil's leading edge. This phenomenon has been observed to have both detrimental (dynamic stall in helicopters) and beneficial (high-lift flight in insects) effects. To predict the formation of LEVs in unsteady flows, a Leading Edge Suction Parameter (LESP) is proposed. This parameter is calculated from inviscid theory and is a measure of the suction at the airfoil's leading edge. It
Bending mode flutter in a transonic linear cascade
Govardhan, Raghuraman; Jutur, Prahallada
2017-11-01
Vibration related issues like flutter pose a serious challenge to aircraft engine designers. The phenomenon has gained relevance for modern engines that employ thin and long fan blade rows to satisfy the growing need for compact and powerful engines. The tip regions of such blade rows operate with transonic relative flow velocities, and are susceptible to bending mode flutter. In such cases, the flow field around individual blades of the cascade is dominated by shock motions generated by the blade motions. In the present work, a new transonic linear cascade facility with the ability to oscillate a blade at realistic reduced frequencies has been developed. The facility operates at a Mach number of 1.3, with the central blade being oscillated in heave corresponding to the bending mode of the rotor. The susceptibility of the blade to undergo flutter at different reduced frequencies is quantified by the cycle-averaged power transfer to the blade calculated using the measured unsteady load on the oscillating blade. These measurements show fluid excitation (flutter) at low reduced frequencies and fluid damping (no flutter) at higher reduced frequencies. Simultaneous measurements of the unsteady shock motions are done with high speed shadowgraphy to elucidate the differences in shock motions between the excitation and damping cases.
Pressure Distribution Over Airfoils with Fowler Flaps
Wenzinger, Carl J; Anderson, Walter B
1938-01-01
Report presents the results of tests made of a Clark y airfoil with a Clark y Fowler flap and of an NACA 23012 airfoil with NACA Fowler flaps. Some of the tests were made in the 7 by 10-foot wind tunnel and others in the 5-foot vertical wind tunnel. The pressures were measured on the upper and lower surfaces at one chord section both on the main airfoils and on the flaps for several angles of attack with the flaps located at the maximum-lift settings. A test installation was used in which the model was mounted in the wind tunnel between large end planes so that two-dimensional flow was approximated. The data are given in the form of pressure-distribution diagrams and as plots of calculated coefficients for the airfoil-and-flap combinations and for the flaps alone.
A finite element solution of transonic flow
Tatum, K. E.
1978-01-01
The use of finite elements is explored in a field in which its use has previously not been deemed very feasible, that of transonic flow. The specific problem chosen is that of steady small-disturbance transonic flow. The nonlinear equations are formulated with an artificial viscosity term added to yield the proper domain of dependence and directional bias in supersonic regions and across imbedded shock waves. Justification is given for the problem and means of solution chosen, and the potential advantages of the finite element procedure over standard finite difference procedures are discussed. Several possible improvements on the method as presently derived are stated. Computational mesh requirements and certain mesh variations are described. Some results equivalent to finite difference calculations are given as a sample solution.
Development of a fast shape memory alloy based actuator for morphing airfoils
Lara-Quintanilla, A.
2016-01-01
The design of aerodynamic airfoils are optimized for certain conditions. For instance, the shape of the wings of fixed-wing aircrafts are designed and optimized for a certain flight condition (in terms of altitude, speed, aircraft weight, etc.). However, these flight conditions vary significantly
Magnus effects on spinning transonic missiles
Seginer, A.; Rosenwasser, I.
1983-01-01
Magnus forces and moments were measured on a basic-finner model spinning in transonic flow. Spin was induced by canted fins or by full-span or semi-span, outboard and inboard roll controls. Magnus force and moment reversals were caused by Mach number, reduced spin rate, and angle of attack variations. Magnus center of pressure was found to be independent of the angle of attack but varied with the Mach number and model configuration or reduced spin rate.
Flow past a self-oscillating airfoil with two degrees of freedom: measurements and simulations
Directory of Open Access Journals (Sweden)
Šidlof Petr
2014-03-01
Full Text Available The paper focuses on investigation of the unsteady subsonic airflow past an elastically supported airfoil for subcritical flow velocities and during the onset of the flutter instability. A physical model of the NACA0015 airfoil has been designed and manufactured, allowing motion with two degrees of freedom: pitching (rotation about the elastic axis and plunging (vertical motion. The structural mass and stiffness matrix can be tuned to certain extent, so that the natural frequencies of the two modes approach as needed. The model was placed in the measuring section of the wind tunnel in the aerodynamic laboratory of the Institute of Thermomechanics in Nový Knín, and subjected to low Mach number airflow up to the flow velocities when self-oscillation reach amplitudes dangerous for the structural integrity of the model. The motion of the airfoil was registered by a high-speed camera, with synchronous measurement of the mechanic vibration and discrete pressure sensors on the surface of the airfoil. The results of the measurements are presented together with numerical simulation results, based on a finite volume CFD model of airflow past a vibrating airfoil.
Experimental Study of the Power Profile Airfoil Equipped with Plasma Flow Control
Daniel, Libin; Jacob, Jamey
2013-11-01
This presentation discusses results from an experimental study of the power profile airfoil at low Reynolds number. The power profile airfoil was developed by AMO Smith and consists of a blunt trailing edge shape with two wall jets near the trailing edge. The replacement of streamlining with properly designed blowing is used to prevent flow separation and additionally offers potential applications as a powered high-lift system, propulsive system, or low inertia control device. The 2D wind-tunnel model consists of the 22.5% thick power profile airfoil equipped with a movable trailing edge plug to direct flow along the trailing edge streamline. Compressed air was passed into the model via a plenum with flow conditioning devices to create pressure backdrop to allow uniform blowing at the trailing edge. The effects of varying jet momentum coefficient and trailing edge positioning on the aerodynamic characteristics are observed with both wake surveys and PIV. The impact of plasma synthetic jet actuators (PSJA) placed along the trailing edge of the power profile airfoil is also discussed. PSJA operation is compared to the baseline power profile airfoil both alone and working with the blowing to provide additional control authority.
METHOD FOR NUMERICAL MODELING OF UNSTEADY SEPARATED FLOW AROUND AIRFOILS MOVING CLOSE TO FLAT SCREEN
Directory of Open Access Journals (Sweden)
V. Pogrebnaya Tamara
2017-01-01
Full Text Available In this article an attempt is made to explain the nature of differences in measurements of forces and moments, which influence an aircraft at take-off and landing when testing on different types of stands. An algorithm for numerical simulation of unsteady separated flow around airfoil is given. The algorithm is based on the combination of discrete vortex method and turbulent boundary layer equations. An unsteady flow separation modeling has been used. At each interval vortex method was used to calculate the potential flow around airfoils located near a screen. Calculated pressures and velocities were then used in boundary layer calculations to determine flow separation points and separated vortex in- tensities. After that calculation were made to determine free vortex positions to next time step and the process was fulfilled for next time step. The proposed algorithm allows using numeric visualization to understand physical picture of flow around airfoil moving close to screen. Three different ways of flow modeling (mirror method, fixed or movable screens were tested. In each case the flow separation process, which determines pressure distribution over airfoil surface and influ- ences aerodynamic performance, was viewed. The results of the calculations showed that at low atitudes of airfoil over screen mirror method over predicts lift force compared with movable screen, while fixed screen under predicts it. The data obtained can be used when designing equipment for testing in wind tunnels.
Model measurements in the cryogenic National Transonic Facility - An overview
Holmes, H. K.
1985-01-01
In the operation of the National Transonic Facility (NTF) higher Reynolds numbers are obtained on the basis of a utilization of low operational temperatures and high pressures. Liquid nitrogen is used as cryogenic medium, and temperatures in the range from -320 F to 160 F can be employed. A maximum pressure of 130 psi is specified, while the NTF design parameter for the Reynolds number is 120,000,000. In view of the new requirements regarding the measurement systems, major developments had to be undertaken in virtually all wind tunnel measurement areas and, in addition, some new measurement systems were needed. Attention is given to force measurement, pressure measurement, model attitude, model deformation, and the data system.
Realizing compact system for Schlieren visualization of transonic flows
International Nuclear Information System (INIS)
Bassi, F.; Osnaghi, C.; Savini, M.
1992-01-01
This paper describes a Schlieren system for flow visualization thought to overcome most of the drawbacks connected with the Z-configuration usually adopted: lack of transportability, wide overall dimensions, long and difficult alignment of the components, restricted sensitivity. The system, originally designed and realized, is of the double-pass type with all the emitting and receiving components mounted on an optical table and aligned during the initial assembly. The paper shows both monochromatic (Helium-Neon laser) and white-light (Mercury or Xenon arc lamps) visualizations of flames and transonic flows in calibrating nozzles and in turbine cascades in a wind tunnel. The accuracy and versatility displayed make this Schlieren system easily usable in fluid dynamics researches
Milholen, William E., II; Jones, Gregory S.; Chan, David T.; Goodliff, Scott L.; Anders, Scott G.; Melton, Latunia P.; Carter, Melissa B.; Allan, Brian G.; Capone, Francis J.
2013-01-01
A second wind tunnel test of the FAST-MAC circulation control model was recently completed in the National Transonic Facility at the NASA Langley Research Center. The model was equipped with four onboard flow control valves allowing independent control of the circulation control plenums, which were directed over a 15% chord simple-hinged flap. The model was configured for low-speed high-lift testing with flap deflections of 30 and 60 degrees, along with the transonic cruise configuration with zero degree flap deflection. Testing was again conducted over a wide range of Mach numbers up to 0.88, and Reynolds numbers up to 30 million based on the mean chord. The first wind tunnel test had poor transonic force and moment data repeatability at mild cryogenic conditions due to inadequate thermal conditioning of the balance. The second test demonstrated that an improvement to the balance heating system significantly improved the transonic data repeatability, but also indicated further improvements are still needed. The low-speed highlift performance of the model was improved by testing various blowing slot heights, and the circulation control was again demonstrated to be effective in re-attaching the flow over the wing at off-design transonic conditions. A new tailored spanwise blowing technique was also demonstrated to be effective at transonic conditions with the benefit of reduced mass flow requirements.
2-D and 3-D CFD Investigation of NREL S826 Airfoil at Low Reynolds Numbers
International Nuclear Information System (INIS)
Cakmakcioglu, S C; Sert, I O; Tugluk, O; Sezer-Uzol, N
2014-01-01
In this study CFD investigation of flow over the NREL S826 airfoil is performed. NREL S826 airfoil was designed for HAWTs of 10-15 meter diameters. However, it is used in the NTNU wind turbine rotor model and low Reynolds number flow characteristics become important in the validations with the test cases of this rotor model. The airfoil CFD simulations are carried out in 2-D and 3-D computational domains. The k-rn SST turbulence model with Langtry-Menter (γ-Re θ ) transition prediction model for turbulence closure is used in the calculations. The Delayed DES is also performed in the stall region for comparisons. The results are compared with the available METUWIND experimental data, and are shown to be in fair agreement. It is observed that 3-D CFD analysis provides increased accuracy at increased computational cost
Multi-element airfoil optimization for maximum lift at high Reynolds numbers
Valarezo, Walter O.; Dominik, Chet J.; Mcghee, Robert J.; Goodman, Wesley L.; Paschal, Keith B.
1991-01-01
An experimental study has been performed to assess the maximum lift capability of a supercritical multielement airfoil representative of an advanced transport aircraft wing. The airfoil model was designed with a leading-edge slat and single or two-segment trailing-edge flaps. Optimization work was performed at various slat/flap deflections as well as gap/overhang positions. Landing configurations and the attainment of maximum lift coefficients of 4.5 with single-element flaps and 5.0 with two-segment flaps was emphasized. Test results showed a relatively linear variation of the optimum gap/overhang positioning of the slat versus slat deflection, considerable differences in optimum rigging between single and double segment flaps, and large Reynolds number effects on multielement airfoil optimization.
Timmer, W.A.
2009-01-01
This paper investigates the NACA 63 and 64 6-digit series of airfoils tested in the NACA LTPT in view to verify the RFOIL calculated airfoil characteristics for high Reynolds numbers. Some anomalies in the zero-lift angles of 15% and 18% thick airfoils from these series are identified, both in the
Experimental parameter study for passive vortex generators on a 30% thick airfoil
Baldacchino, D.; Simao Ferreira, C.; De Tavernier, D.A.M.; Timmer, W.A.; van Bussel, G.J.W.
2018-01-01
Passive vane-type vortex generators (VGs) are commonly used on wind turbine blades to mitigate the effects of flow separation. However, significant uncertainty surrounds VG design guidelines. Understanding the influence of VG parameters on airfoil performance requires a systematic approach
Sydnor, George H.; Bhatia, Ram; Krattiger, Hansueli; Mylius, Justus; Schafer, D.
2012-01-01
In September 1995, a project was initiated to replace the existing drive line at NASA's most unique transonic wind tunnel, the National Transonic Facility (NTF), with a single 101 MW synchronous motor driven by a Load Commutated Inverter (LCI). This Adjustable Speed Drive (ASD) system also included a custom four-winding transformer, harmonic filter, exciter, switch gear, control system, and feeder cable. The complete system requirements and design details have previously been presented and published [1], as well as the commissioning and acceptance test results [2]. The NTF was returned to service in December 1997 with the new drive system powering the fan. Today, this installation still represents the world s largest horizontal single motor/drive combination. This paper describes some significant events that occurred with the drive system during the first 15 years of service. These noteworthy issues are analyzed and root causes presented. Improvements that have substantially increased the long term viability of the system are given.
New airfoil sections for straight bladed turbine
International Nuclear Information System (INIS)
Boumaza, B.
1987-07-01
A theoretical investigation of aerodynamic performance for vertical axis Darrieus wind turbine with new airfoils sections is carried out. The blade section aerodynamics characteristics are determined from turbomachines cascade model. The model is also adapted to the vertical Darrieus turbine for the performance prediction of the machine. In order to choose appropriate value of zero-lift-drag coefficient in calculation, an analytical expression is introduced as function of chord-radius ratio and Reynolds numbers. New airfoils sections are proposed and analyzed for straight-bladed turbine
EUDP Project: Low Noise Airfoil - Final Report
DEFF Research Database (Denmark)
. In particular, the so-called TNO trailing edge noise model could be significantly improved by introducing turbulence anisotropy in its formulation, as well as the influence of the boundary layer mean pressure gradient. This two characteristics are inherent to airfoil flows but were neglected in the original...
Investigation of Unsteady Flow Behavior in Transonic Compressor Rotors with LES and PIV Measurements
Hah, Chunill; Voges, Melanie; Mueller, Martin; Schiffer, Heinz-Peter
2009-01-01
In the present study, unsteady flow behavior in a modern transonic axial compressor rotor is studied in detail with large eddy simulation (LES) and particle image velocimetry (PIV). The main purpose of the study is to advance the current understanding of the flow field near the blade tip in an axial transonic compressor rotor near the stall and peak-efficiency conditions. Flow interaction between the tip leakage vortex and the passage shock is inherently unsteady in a transonic compressor. Casing-mounted unsteady pressure transducers have been widely applied to investigate steady and unsteady flow behavior near the casing. Although many aspects of flow have been revealed, flow structures below the casing cannot be studied with casing-mounted pressure transducers. In the present study, unsteady velocity fields are measured with a PIV system and the measured unsteady flow fields are compared with LES simulations. The currently applied PIV measurements indicate that the flow near the tip region is not steady even at the design condition. This self-induced unsteadiness increases significantly as the compressor rotor operates near the stall condition. Measured data from PIV show that the tip clearance vortex oscillates substantially near stall. The calculated unsteady characteristics of the flow from LES agree well with the PIV measurements. Calculated unsteady flow fields show that the formation of the tip clearance vortex is intermittent and the concept of vortex breakdown from steady flow analysis does not seem to apply in the current flow field. Fluid with low momentum near the pressure side of the blade close to the leading edge periodically spills over into the adjacent blade passage. The present study indicates that stall inception is heavily dependent on unsteady behavior of the flow field near the leading edge of the blade tip section for the present transonic compressor rotor.
Historical Review of Uncommanded Lateral-Directional Motions at Transonic Conditions
Chambers, Joseph R.; Hall, Robert M.
2003-01-01
This paper presents the results of a survey of past experiences with uncommanded lateral-directional motions at transonic speeds during specific military aircraft programs. The effort was undertaken to provide qualitative and quantitative information on past airplane programs that might be of use to the participants in the joint NASA/Navy/Air Force Abrupt Wing Stall (AWS) Program. The AWS Program was initiated because of the experiences of the F/A-l8E/F development program, during which unexpected, severe wing-drop motions were encountered by preproduction aircraft at transonic conditions. These motions were judged to be significantly degrading to the primary mission requirements of the aircraft. Although the problem was subsequently solved for the production version of the F/A-l8E/F, a high-level review panel emphasized the poor understanding of such phenomena and issued a strong recommendation to: "Initiate a national research effort to thoroughly and systematically study the wing drop phenomena." A comprehensive, cooperative NASA/Navy/Air Force AWS Program was designed to respond to provide the required technology requirements. As part of the AWS Program, a work element was directed at a historical review of wing-drop experiences in past aircraft development programs at high subsonic and transonic speeds. In particular, information was requested regarding: specific aircraft configurations that exhibited uncommanded motions and the nature of the motions; geometric characteristics of the air- planes; flight conditions involved in occurrences; relevant data, including wind-tunnel, computational, and flight sources; figures of merit used for analyses; and approaches used to alleviate the problem. An attempt was also made to summarize some of the more important lessons learned from past experiences, and to recommend specific research efforts. In addition to providing technical information to assist the AWS research objectives, the study produced fundamental
Historical Review of Uncommanded Lateral-Directional Motions At Transonic Conditions (Invited)
Chambers, Joseph R.; Hall, Robert M.
2003-01-01
This paper presents the results of a survey of past experiences with uncommanded lateral-directional motions at transonic speeds during specific military aircraft programs. The effort was undertaken to provide qualitative and quantitative information on past airplane programs that might be of use to the participants in the joint NASA/Navy/Air Force Abrupt Wing Stall (AWS) Program. The AWS Program was initiated because of the experiences of the F/A-18E/F development program, during which unexpected, severe wing-drop motions were encountered by preproduction aircraft at transonic conditions. These motions were judged to be significantly degrading to the primary mission requirements of the aircraft. Although the problem was subsequently solved for the production version of the F/A-l8E/F, a high-level review panel emphasized the poor understanding of such phenomena and issued a strong recommendation to: Initiate a national research effort to thoroughly and systematically study the wing drop phenomena. A comprehensive, cooperative NASA/Navy/Air Force AWS Program was designed to respond to provide the required technology requirements. As part of the AWS Program, a work element was directed at a historical review of wing-drop experiences in past aircraft development programs at high subsonic and transonic speeds. In particular, information was requested regarding: specific aircraft configurations that exhibited uncommanded motions and the nature of the motions; geometric characteristics of the air- planes; flight conditions involved in occurrences; relevant data, including wind-tunnel, computational, and flight sources; figures of merit used for analyses; and approaches used to alleviate the problem. An attempt was also made to summarize some of the more important lessons learned from past experiences, and to recommend specific research efforts. In addition to providing technical information to assist the AWS research objectives, the study produced fundamental information
Ivanco, Thomas G.
2013-01-01
NASA Langley Research Center's Transonic Dynamics Tunnel (TDT) is the world's most capable aeroelastic test facility. Its large size, transonic speed range, variable pressure capability, and use of either air or R-134a heavy gas as a test medium enable unparalleled manipulation of flow-dependent scaling quantities. Matching these scaling quantities enables dynamic similitude of a full-scale vehicle with a sub-scale model, a requirement for proper characterization of any dynamic phenomenon, and many static elastic phenomena. Select scaling parameters are presented in order to quantify the scaling advantages of TDT and the consequence of testing in other facilities. In addition to dynamic testing, the TDT is uniquely well-suited for high risk testing or for those tests that require unusual model mount or support systems. Examples of recently conducted dynamic tests requiring unusual model support are presented. In addition to its unique dynamic test capabilities, the TDT is also evaluated in its capability to conduct aerodynamic performance tests as a result of its flow quality. Results of flow quality studies and a comparison to a many other transonic facilities are presented. Finally, the ability of the TDT to support future NASA research thrusts and likely vehicle designs is discussed.
International Nuclear Information System (INIS)
Thiery, Mylene; Coustols, Eric
2006-01-01
The present study deals with recent numerical results from on-going research conducted at ONERA/DMAE regarding the prediction of transonic flows, for which shock wave/boundary layer interaction is important. When this interaction is strong enough (M ≥ 1.3), shock induced oscillations (SIO) appear at the suction side of the airfoil and lead to the formation of unsteady separated areas. The main issue is then to perform unsteady computations applying appropriate turbulence modelling and relevant boundary conditions with respect to the test case. Computations were performed with the ONERA elsA software and the URANS-type approach, closure relationships being achieved from transport-equation models. Applications are provided for the OAT15A airfoil data base, well documented for unsteady CFD validation (mean and r.m.s. pressure, phase-averaged LDA data, ...). In this paper, the capabilities of turbulence models are evaluated with two 2D URANS strategies, under free-stream or confined conditions. The latter takes into account the adaptive upper and lower wind-tunnel walls. A complete 3D URANS simulation was then performed to demonstrate the real impact of all lateral wind-tunnel walls on such a flow
Transonic Wing Shape Optimization Using a Genetic Algorithm
Holst, Terry L.; Pulliam, Thomas H.; Kwak, Dochan (Technical Monitor)
2002-01-01
A method for aerodynamic shape optimization based on a genetic algorithm approach is demonstrated. The algorithm is coupled with a transonic full potential flow solver and is used to optimize the flow about transonic wings including multi-objective solutions that lead to the generation of pareto fronts. The results indicate that the genetic algorithm is easy to implement, flexible in application and extremely reliable.
Remark on numerical simulation of 2D unsteady transonic flows
Foŕt, J.; Hülek, T.; Kozel, K.; Vavrincová, M.
The work deals with three numerical methods solving the system of Euler or Navier-Stokes equations. Mac Cormack cell centered and Ni cell vertex finite volume schemes were used for simulation of inviscid unsteady solution of transonic flows through a 2D cascade. Unsteady motion is caused by a periodic change of downstream pressure. The Runge-Kutta multistage cell centered finite volume scheme has been used for viscous laminar steady and unsteady transonic flows over NACA 0012.
An approach for aerodynamic optimization of transonic fan blades
Khelghatibana, Maryam
Aerodynamic design optimization of transonic fan blades is a highly challenging problem due to the complexity of flow field inside the fan, the conflicting design requirements and the high-dimensional design space. In order to address all these challenges, an aerodynamic design optimization method is developed in this study. This method automates the design process by integrating a geometrical parameterization method, a CFD solver and numerical optimization methods that can be applied to both single and multi-point optimization design problems. A multi-level blade parameterization is employed to modify the blade geometry. Numerical analyses are performed by solving 3D RANS equations combined with SST turbulence model. Genetic algorithms and hybrid optimization methods are applied to solve the optimization problem. In order to verify the effectiveness and feasibility of the optimization method, a singlepoint optimization problem aiming to maximize design efficiency is formulated and applied to redesign a test case. However, transonic fan blade design is inherently a multi-faceted problem that deals with several objectives such as efficiency, stall margin, and choke margin. The proposed multi-point optimization method in the current study is formulated as a bi-objective problem to maximize design and near-stall efficiencies while maintaining the required design pressure ratio. Enhancing these objectives significantly deteriorate the choke margin, specifically at high rotational speeds. Therefore, another constraint is embedded in the optimization problem in order to prevent the reduction of choke margin at high speeds. Since capturing stall inception is numerically very expensive, stall margin has not been considered as an objective in the problem statement. However, improving near-stall efficiency results in a better performance at stall condition, which could enhance the stall margin. An investigation is therefore performed on the Pareto-optimal solutions to
Demonstration of PIV in a Transonic Compressor
Wernet, Mark P.
1998-01-01
Particle Imaging Velocimetry (PIV) is a powerful measurement technique which can be used as an alternative or complementary approach to Laser Doppler Velocimetry (LDV) in a wide range of research applications. PIV data are measured simultaneously at multiple points in space, which enables the investigation of the non-stationary spatial structures typically encountered in turbomachinery. Many of the same issues encountered in the application of LDV techniques to rotating machinery apply in the application of PIV. Preliminary results from the successful application of the standard 2-D PIV technique to a transonic axial compressor are presented. The lessons learned from the application of the 2-D PIV technique will serve as the basis for applying 3-component PIV techniques to turbomachinery.
Analysis of three-dimensional transonic compressors
Bourgeade, A.
1984-01-01
A method for computing the three-dimensional transonic flow around the blades of a compressor or of a propeller is given. The method is based on the use of the velocity potential, on the hypothesis that the flow is inviscid, irrotational and isentropic. The equation of the potential is solved in a transformed space such that the surface of the blade is mapped into a plane where the periodicity is implicit. This equation is in a nonconservative form and is solved with the help of a finite difference method using artificial time. A computer code is provided and some sample results are given in order to demonstrate the influence of three-dimensional effects and the blade's rotation.
Numerical simulation of transonic flows in diffusers
Liou, M.-S.; Coakley, T. J.; Bergmann, M. Y.
1981-01-01
Numerical simulations were made of two-dimensional transonic flows in diffusers, including flow separation induced by a shock or adverse pressure gradient. The mass-averaged, time-dependent, compressible Navier-Stokes equations, simplified by the thin-layer approximation, were solved using MacCormack's hybrid method. The eddy-viscosity formulation was described by the Wilcox-Rubesin's two-equation, k-omega model. Detailed comparison of the computed results with measurements showed good agreement in all cases, including one with massive separation induced by a strong shock. The computation correctly predicted the details of a distinct lambda shock pattern, closely duplicating the configuration observed experimentally in spark-schlieren photographs.
Virtual Shaping of a Two-dimensional NACA 0015 Airfoil Using Synthetic Jet Actuator
Chen, Fang-Jenq; Beeler, George B.
2002-01-01
The Aircraft Morphing Program at NASA Langley envisions an aircraft without conventional control surfaces. Instead of moving control surfaces, the vehicle control systems may be implemented with a combination of propulsive forces, micro surface effectors, and fluidic devices dynamically operated by an intelligent flight control system to provide aircraft maneuverability over each mission segment. As a part of this program, a two-dimensional NACA 0015 airfoil model was designed to test mild maneuvering capability of synthetic jets in a subsonic wind tunnel. The objective of the experiments is to assess the applicability of using unsteady suction and blowing to alter the aerodynamic shape of an airfoil with a purpose to enhance lift and/or to reduce drag. Synthetic jet actuation at different chordwise locations, different forcing frequencies and amplitudes, under different freestream velocities are investigated. The effect of virtual shape change is indicated by a localized increase of surface pressure in the neighborhood of synthetic jet actuation. That causes a negative lift to the airfoil with an upper surface actuation. When actuation is applied near the airfoil leading edge, it appears that the stagnation line is shifted inducing an effect similar to that caused by a small angle of attack to produce an overall lift change.
Vortex-Induced Vibration of an Airfoil Used in Vertical-Axis Wind Turbines
Benner, Bridget; Carlson, Daniel; Seyed-Aghazadeh, Banafsheh; Modarres-Sadeghi, Yahya
2017-11-01
In Vertical-axis wind turbines (VAWTs), when the blades are placed at high angles of attack with respect to the incoming flow, they could experience flow-induced oscillations. A series of experiments in a re-circulating water tunnel was conducted to study the possible Vortex-Induced Vibration (VIV) of a fully-submerged, flexibly-mounted NACA 0021 airfoil, which is used in some designs of VAWTs. The airfoil was free to oscillate in the crossflow direction, and the tests were conducted in a Reynolds number range of 600
CFD study on NACA 4415 airfoil implementing spherical and sinusoidal Tubercle Leading Edge.
Directory of Open Access Journals (Sweden)
S M A Aftab
Full Text Available The Humpback whale tubercles have been studied for more than a decade. Tubercle Leading Edge (TLE effectively reduces the separation bubble size and helps in delaying stall. They are very effective in case of low Reynolds number flows. The current Computational Fluid Dynamics (CFD study is on NACA 4415 airfoil, at a Reynolds number 120,000. Two TLE shapes are tested on NACA 4415 airfoil. The tubercle designs implemented on the airfoil are sinusoidal and spherical. A parametric study is also carried out considering three amplitudes (0.025c, 0.05c and 0.075c, the wavelength (0.25c is fixed. Structured mesh is utilized to generate grid and Transition SST turbulence model is used to capture the flow physics. Results clearly show spherical tubercles outperform sinusoidal tubercles. Furthermore experimental study considering spherical TLE is carried out at Reynolds number 200,000. The experimental results show that spherical TLE improve performance compared to clean airfoil.
Turbine airfoil having near-wall cooling insert
Energy Technology Data Exchange (ETDEWEB)
Martin, Jr., Nicholas F.; Wiebe, David J.
2017-09-12
A turbine airfoil is provided with at least one insert positioned in a cavity in an airfoil interior. The insert extends along a span-wise extent of the turbine airfoil and includes first and second opposite faces. A first near-wall cooling channel is defined between the first face and a pressure sidewall of an airfoil outer wall. A second near-wall cooling channel is defined between the second face and a suction sidewall of the airfoil outer wall. The insert is configured to occupy an inactive volume in the airfoil interior so as to displace a coolant flow in the cavity toward the first and second near-wall cooling channels. A locating feature engages the insert with the outer wall for supporting the insert in position. The locating feature is configured to control flow of the coolant through the first or second near-wall cooling channel.
Aerodynamic and aeroacoustic performance of airfoils with morphing structures
Ai, Qing; Azarpeyvand, Mahdi; Lachenal, Xavier; Weaver, Paul M.
2016-01-01
Aerodynamic and aeroacoustic performance of airfoils fitted with morphing trailing edges are investigated using a coupled structure/fluid/noise model. The control of the flow over the surface of an airfoil using shape optimization techniques can significantly improve the load distribution along the chord and span lengths whilst minimising noise generation. In this study, a NACA 63-418 airfoil is fitted with a morphing flap and various morphing profiles are considered with two features that di...
Damping element for reducing the vibration of an airfoil
Campbell, Christian X; Marra, John J
2013-11-12
An airfoil (10) is provided with a tip (12) having an opening (14) to a center channel (24). A damping element (16) is inserted within the opening of the center channel, to reduce an induced vibration of the airfoil. The mass of the damping element, a spring constant of the damping element within the center channel, and/or a mounting location (58) of the damping element within the center channel may be adjustably varied, to shift a resonance frequency of the airfoil outside a natural operating frequency of the airfoil.
Profile catalogue for airfoil sections based on 3D computations
DEFF Research Database (Denmark)
Bertagnolio, F.; Sørensen, Niels N.; Johansen, Jeppe
2006-01-01
This report is a continuation of the Wind Turbine Airfoil Catalogue [1] which objective was, firstly to provide a database of aerodynamic characteristics for a wide range of airfoil profiles aimed at wind turbine applications, and secondly to test thetwo-dimensional Navier-Stokes solver EllipSys2D...... and the actual fluid flow, and thereby the incorrect prediction of airfoil characteristics. In addition, other features of the flow solver, such astransition and turbulence modelling, and their influence onto the numerical results are investigated. Conclusions are drawn regarding the evaluation of airfoil...... aerodynamic characteristics, as well as the use of the Navier-Stokes solver for fluid flowcalculations in general....
Airfoil stall interpreted through linear stability analysis
Busquet, Denis; Juniper, Matthew; Richez, Francois; Marquet, Olivier; Sipp, Denis
2017-11-01
Although airfoil stall has been widely investigated, the origin of this phenomenon, which manifests as a sudden drop of lift, is still not clearly understood. In the specific case of static stall, multiple steady solutions have been identified experimentally and numerically around the stall angle. We are interested here in investigating the stability of these steady solutions so as to first model and then control the dynamics. The study is performed on a 2D helicopter blade airfoil OA209 at low Mach number, M 0.2 and high Reynolds number, Re 1.8 ×106 . Steady RANS computation using a Spalart-Allmaras model is coupled with continuation methods (pseudo-arclength and Newton's method) to obtain steady states for several angles of incidence. The results show one upper branch (high lift), one lower branch (low lift) connected by a middle branch, characterizing an hysteresis phenomenon. A linear stability analysis performed around these equilibrium states highlights a mode responsible for stall, which starts with a low frequency oscillation. A bifurcation scenario is deduced from the behaviour of this mode. To shed light on the nonlinear behavior, a low order nonlinear model is created with the same linear stability behavior as that observed for that airfoil.
Aerodynamic Flow Control of a Maneuvering Airfoil
Brzozowski, Daniel P.; Culp, John; Glezer, Ari
2010-11-01
The unsteady aerodynamic forces and moments on a maneuvering, free-moving airfoil are varied in wind tunnel experiments by controlling vorticity generation/accumulation near the surface using hybrid synthetic jet actuators. The dynamic characteristics of the airfoil that is mounted on a 2-DOF traverse are controlled using position and attitude feedback loops that are actuated by servo motors. Bi-directional changes in the pitching moment are induced using controllable trapped vorticity concentrations on the suction and pressure surfaces near the trailing edge. The dynamic coupling between the actuation and the time-dependent flow field is characterized using simultaneous force and velocity measurements that are taken phase-locked to the commanded actuation waveform. The time scales associated with the actuation process is determined from PIV measurements of vorticity flux downstream of the trailing edge. Circulation time history shows that the entire flow over the airfoil readjusts within about 1.5 TCONV, which is about two orders of magnitude shorter than the characteristic time associated with the controlled maneuver of the wind tunnel model. This illustrates that flow-control actuation can be typically effected on time scales commensurate with the flow's convective time scale, and that the maneuver response is only limited by the inertia of the platform. Supported by AFSOR.
DEFF Research Database (Denmark)
Troldborg, Niels
2005-01-01
A comprehensive computational study, in both steady and unsteady flow conditions, has been carried out to investigate the aerodynamic characteristics of the Risø-B1.18 airfoil equipped with variable trailing edge geometry as produced by a hinged flap. The function of such flaps should...... on the baseline airfoil showed excellent agreement with measurements on the same airfoil with the same specified conditions. Furthermore, a more widespread comparison with an advanced potential theory code is presented. The influence of various key parameters, such as flap shape, flap size and oscillating...... frequencies, was investigated so that an optimum design can be suggested for application with wind turbine blades. It is concluded that a moderately curved flap with flap chord to airfoil curve ratio between 0.05 and 0.10 would be an optimum choice....
Passive Boundary Layer Separation Control on a NACA2415 Airfoil at High Reynolds Numbers
Parikh, Agastya; Hultmark, Marcus
2016-11-01
The design and analysis of a passive flow control system for a NACA2415 airfoil is undertaken. There exists a vast body of knowledge on airfoil boundary layer control with the use of controlled mass flux, but there is little work investigating passive mass flux-based methods. A simple duct system that uses the upper surface pressure gradient to force blowing near the leading edge and suction near the trailing edge is proposed and evaluated. 2D RANS analyses at Rec 1 . 27 ×106 were used to generate potential configurations for experimental tests. Initial computational results suggest drag reductions of approximately 2 - 7 % as well as lift increases of 4 - 5 % at α = 10 .0° and α = 12 .5° . A carbon composite-aluminum structure model that implements the most effective configurations, according to the CFD predictions, has been designed and fabricated. Experiments are being performed to evaluate the CFD results and the feasibility the duct system.
Directory of Open Access Journals (Sweden)
Ziaul Huque
2012-01-01
Full Text Available A Computational Fluid Dynamics (CFD and response surface-based multiobjective design optimization were performed for six different 2D airfoil profiles, and the Pareto optimal front of each airfoil is presented. FLUENT, which is a commercial CFD simulation code, was used to determine the relevant aerodynamic loads. The Lift Coefficient (CL and Drag Coefficient (CD data at a range of 0° to 12° angles of attack (α and at three different Reynolds numbers (Re=68,459, 479, 210, and 958, 422 for all the six airfoils were obtained. Realizable k-ε turbulence model with a second-order upwind solution method was used in the simulations. The standard least square method was used to generate response surface by the statistical code JMP. Elitist Non-dominated Sorting Genetic Algorithm (NSGA-II was used to determine the Pareto optimal set based on the response surfaces. Each Pareto optimal solution represents a different compromise between design objectives. This gives the designer a choice to select a design compromise that best suits the requirements from a set of optimal solutions. The Pareto solution set is presented in the form of a Pareto optimal front.
Identification of dynamic properties of radial air-foil bearings
Arora, Vikas; Arora, V.; van der Hoogt, Peter; Aarts, Ronald G.K.M.; de Boer, Andries
2010-01-01
Air-foil bearings (AFBs) are self acting hydrodynamic bearings made from sheet metal foils comprised of at least two layers. The innermost “top foil” layer traps a gas pressure film that supports a load while the layer or layers underneath provide an elastic foundation. Air-foil bearings are
2015-04-23
CLASSIFICATION OF: The requirement for higher power-to-weight ratios in modern jet engines leads to a reduced number of stages at increased loading...Stage - Final Report Report Title The requirement for higher power-to-weight ratios in modern jet engines leads to a reduced number of stages at...Garth Hobson, PhD, Anthony Gannon, PhD, Scott Drayton, LCDR USN. DESIGN AND TEST OF A TRANSONIC AXIAL SPLITTERED ROTOR, ASME TURBO Expo 2015. 15
Modeling of Airfoil Trailing Edge Flap with Immersed Boundary Method
DEFF Research Database (Denmark)
Zhu, Wei Jun; Shen, Wen Zhong; Sørensen, Jens Nørkær
2011-01-01
The present work considers incompressible flow over a 2D airfoil with a deformable trailing edge. The aerodynamic characteristics of an airfoil with a trailing edge flap is numerically investigated using computational fluid dynamics. A novel hybrid immersed boundary (IB) technique is applied...... to simulate the moving part of the trailing edge. Over the main fixed part of the airfoil the Navier-Stokes (NS) equations are solved using a standard body-fitted finite volume technique whereas the moving trailing edge flap is simulated with the immersed boundary method on a curvilinear mesh. The obtained...... results show that the hybrid approach is an efficient and accurate method for solving turbulent flows past airfoils with a trailing edge flap and flow control using trailing edge flap is an efficient way to regulate the aerodynamic loading on airfoils....
Numerical Optimization of Circulation Control Airfoils.
1980-04-01
the jet height at the slot. 6 COMPUTER PROGRAM The method has been coded in FORTRAN and is known as the CIRCON program. It has been written in an...analysis programs. The program flow chart for such an operation is shown in Figure 5. In so doing, the airfoil analysis program CIRCON is merged with the...4 -4- 944 44 0 0 0o 4) 0 AJj IOC Q) 0M 5. CU -4 10 0 440 10 0 c 4-4 0 0 00 04 z 44 4-0) 00 40 02 C 0 a. c 200 -24 -THEORY ( CIRCON ) -THEORY ( CIRCON
An Investigation of Transonic Resonance in a Mach 2.2 Round Convergent-Divergent Nozzle
Dippold, Vance F., III; Zaman, Khairul B. M. Q.
2015-01-01
Hot-wire and acoustic measurements were taken for a round convergent nozzle and a round convergent-divergent (C-D) nozzle at a jet Mach number of 0.61. The C-D nozzle had a design Mach number of 2.2. Compared to the convergent nozzle jet flow, the Mach 2.2 nozzle jet flow produced excess broadband noise (EBBN). It also produced a transonic resonance tone at 1200 Herz. Computational simulations were performed for both nozzle flows. A steady Reynolds-Averaged Navier-Stokes simulation was performed for the convergent nozzle jet flow. For the Mach 2.2 nozzle flow, a steady RANS simulation, an unsteady RANS (URANS) simulation, and an unsteady Detached Eddy Simulation (DES) were performed. The RANS simulation of the convergent nozzle showed good agreement with the hot-wire velocity and turbulence measurements, though the decay of the potential core was over-predicted. The RANS simulation of the Mach 2.2 nozzle showed poor agreement with the experimental data, and more closely resembled an ideally-expanded jet. The URANS simulation also showed qualitative agreement with the hot-wire data, but predicted a transonic resonance at 1145 Herz. The DES showed good agreement with the hot-wire velocity and turbulence data. The DES also produced a transonic tone at 1135 Herz. The DES solution showed that the destabilization of the shock-induced separation region inside the nozzle produced increased levels of turbulence intensity. This is likely the source of the EBBN.
Unsteady Newton-Busemann flow theory. I - Airfoils
Hui, W. H.; Tobak, M.
1981-01-01
Newtonian flow theory for unsteady flow at very high Mach numbers is completed by the addition of a centrifugal force correction to the impact pressures. The correction term is the unsteady counterpart of Busemann's centrifugal force correction to impact pressures in steady flow. For airfoils of arbitary shape, exact formulas for the unsteady pressure and stiffness and damping-in-pitch derivatives are obtained in closed form, which require only numerical quadratures of terms involving the airfoil shape. They are applicable to airfoils of arbitrary thickness having sharp or blunt leading edges. For wedges and thin airfoils these formulas are greatly simplified, and it is proved that the pitching motions of thin airfoils of convex shape and of wedges of arbitrary thickness are always dynamically stable according to Newton-Busemann theory. Leading-edge bluntness is shown to have a favorable effect on the dynamic stability; on the other hand, airfoils of concave shape tend toward dynamic instability over a range of axis positions if the surface curvature exceeds a certain limit. As a byproduct, it is also shown that a pressure formula recently given by Barron and Mandl for unsteady Newtonian flow over a pitching power-law shaped airfoil is erroneous and that their conclusion regarding the effect of pivot position on the dynamic stability is misleading.
Overview of the Space Launch System Transonic Buffet Environment Test Program
Piatak, David J.; Sekula, Martin K.; Rausch, Russ D.; Florance, James R.; Ivanco, Thomas G.
2015-01-01
Fluctuating aerodynamic loads are a significant concern for the structural design of a launch vehicle, particularly while traversing the transonic flight environment. At these trajectory conditions, unsteady aerodynamic pressures can excite the vehicle dynamic modes of vibration and result in high structural bending moments and vibratory environments. To ensure that vehicle structural components and subsystems possess adequate strength, stress, and fatigue margins in the presence of buffet and other environments, buffet forcing functions are required to conduct the coupled load analysis of the launch vehicle. The accepted method to obtain these buffet forcing functions is to perform wind-tunnel testing of a rigid model that is heavily instrumented with unsteady pressure transducers designed to measure the buffet environment within the desired frequency range. Two wind-tunnel tests of a 3 percent scale rigid buffet model have been conducted at the Langley Research Center Transonic Dynamics Tunnel (TDT) as part of the Space Launch System (SLS) buffet test program. The SLS buffet models have been instrumented with as many as 472 unsteady pressure transducers to resolve the buffet forcing functions of this multi-body configuration through integration of the individual pressure time histories. This paper will discuss test program development, instrumentation, data acquisition, test implementation, data analysis techniques, and several methods explored to mitigate high buffet environment encountered during the test program. Preliminary buffet environments will be presented and compared using normalized sectional buffet forcing function root-meansquared levels along the vehicle centerline.
Second-order subsonic airfoil theory including edge effects
Van Dyke, Milton D
1956-01-01
Several recent advances in plane subsonic flow theory are combined into a unified second-order theory for airfoil sections of arbitrary shape. The solution is reached in three steps: the incompressible result is found by integration, it is converted into the corresponding subsonic compressible result by means of the second-order compressibility rule, and it is rendered uniformly valid near stagnation points by further rules. Solutions for a number of airfoils are given and are compared with the results of other theories and of experiment. A straight-forward computing scheme is outlined for calculating the surface velocities and pressures on any airfoil at any angle of attack
Geometrical effects on the airfoil flow separation and transition
Zhang, Wei
2015-04-25
We present results from direct numerical simulations (DNS) of incompressible flow over two airfoils, NACA-4412 and NACA-0012-64, to investigate the effects of the airfoil geometry on the flow separation and transition patterns at Re=104 and 10 degrees incidence. The two chosen airfoils are geometrically similar except for maximum camber (respectively 4%C and 0 with C the chord length), which results in a larger projection area with respect to the incoming flow for the NACA-4412 airfoil, and a larger leeward surface curvature at the leading edge for the NACA-0012-64 airfoil. The governing equations are discretized using an energy conservative fourth-order spatial discretization scheme. An assessment on the two-point correlation indicates that a spanwise domain size of 0.8C is sufficiently large for the present simulations. We discuss flow separation at the airfoil leading edge, transition of the separated shear layer to three-dimensional flow and subsequently to turbulence. Numerical results reveal a stronger adverse pressure gradient field in the leading edge region of the NACA-0012-64 airfoil due to the rapidly varying surface curvature. As a result, the flow experiences detachment at x/C=0.08, and the separated shear layer transition via Kelvin-Helmholtz mechanism occurs at x/C=0.29 with fully developed turbulent flow around x/C=0.80. These flow development phases are delayed to occur at much downstream positions, respectively, observed around x/C=0.25, 0.71 and 1.15 for the NACA-4412 airfoil. The turbulent intensity, measured by the turbulent fluctuations and turbulent Reynolds stresses, are much larger for NACA-0012-64 from the transition onset until the airfoil trailing edge, while turbulence develops significantly downstream of the trailing edge for the NACA-4412 airfoil. For both airfoils, our DNS results indicate that the mean Reynolds stress u\\'u\\'/U02 reaches its maximum value at a distance from the surface approximately equal to the displacement
Experiences with optimizing airfoil shapes for maximum lift over drag
Doria, Michael L.
1991-01-01
The goal was to find airfoil shapes which maximize the ratio of lift over drag for given flow conditions. For a fixed Mach number, Reynolds number, and angle of attack, the lift and drag depend only on the airfoil shape. This then becomes a problem in optimization: find the shape which leads to a maximum value of lift over drag. The optimization was carried out using a self contained computer code for finding the minimum of a function subject to constraints. To find the lift and drag for each airfoil shape, a flow solution has to be obtained. This was done using a two dimensional Navier-Stokes code.
DETERMINATION OF PARASITE DRAG FOR COMMERCIAL AIRPLANE AT TRANSONIC SPEEDS
V. I. Shevyakov
2014-01-01
It considers the task of determination of parasite drag for subsonic commercial airplane on the base of theoretical calculation procedures. It contains the procedure for clarification of empirical approach to calculation of parasite drag at high speeds. It provides examples of calculation of parasite drag at transonic speeds as a result of step behind the slat on the wing of commercial airplane.
Řidký, V.; Šidlof, P.; Vlček, V.
2013-04-01
The work is devoted to comparing measured data with the results of numerical simulations. As mathematical model was used mathematical model whitout turbulence for incompressible flow In the experiment was observed the behavior of designed NACA0015 airfoil in airflow. For the numerical solution was used OpenFOAM computational package, this is open-source software based on finite volume method. In the numerical solution is prescribed displacement of the airfoil, which corresponds to the experiment. The velocity at a point close to the airfoil surface is compared with the experimental data obtained from interferographic measurements of the velocity field. Numerical solution is computed on a 3D mesh composed of about 1 million ortogonal hexahedron elements. The time step is limited by the Courant number. Parallel computations are run on supercomputers of the CIV at Technical University in Prague (HAL and FOX) and on a computer cluster of the Faculty of Mechatronics of Liberec (HYDRA). Run time is fixed at five periods, the results from the fifth periods and average value for all periods are then be compared with experiment.
Flow Quality Measurements in the NASA Ames Upgraded 11-by 11-Foot Transonic Wind Tunnel
Amaya, Max A.; Murthy, Sreedhara V.; George, M. W. (Technical Monitor)
2000-01-01
Among the many upgrades designed and implemented in the NASA Ames 11-by 11-Foot Transonic Wind Tunnel over the past few years, several directly affect flow quality in the test section: a turbulence reduction system with a honeycomb and two screens, a flow smoothing system in the back leg diffusers, an improved drive motor control system, and a full replacement set of composite blades for the compressor. Prior to the shut-down of the tunnel for construction activities, an 8-foot span rake populated with flow instrumentation was traversed in the test section to fully document the flow quality and establish a baseline against which the upgrades could be characterized. A similar set of measurements was performed during the recent integrated system test trials, but the scope was somewhat limited in accordance with the primary objective of such tests, namely to return the tunnel to a fully operational status. These measurements clearly revealed substantial improvements in flow angularity and significant reductions in turbulence level for both full-span and semi-span testing configurations, thus making the flow quality of the tunnel one of the best among existing transonic facilities.
Mechanisms of Sweep on the Performance of Transonic Centrifugal Compressor Impellers
Directory of Open Access Journals (Sweden)
Xiao He
2017-10-01
Full Text Available Transonic centrifugal compressors with high performance are required in the oil and gas industries, modern gas turbine engines, and turbochargers. The sweep of the blades is one of the crucial features that have a significant influence on their performance. This paper numerically investigates mechanisms by which sweep affects the performance of a transonic impeller with twin splitters. Sweep is defined as scaling up or down the shroud chord, and the variation range of the sweep angle has been chosen from −25 to +25°. In the current case, results show that the variation of choke mass flow rate, pressure ratio, and efficiency value is around 1%. If the centrifugal compressor has a higher pressure ratio or a higher front loading, the sweep effect on compressor performance will be even stronger. The essential aerodynamic effect of sweep is the spanwise redistribution of the front loading, resulting in effects on the shock structure, the tip leakage vortex, and the flow separation. On the shroud section, forward sweep restricts the front loading, the shock strength, and the tip leakage vortex, which reduces the loss near the casing. On the hub section, aft sweep suppresses the front loading and the flow separation, which reduces the loss near the hub. It is the delicate balance between controlling the loss near the hub and the loss near the casing that determines the optimal sweep angle design.
Stability investigation of an airfoil section with active flap control
DEFF Research Database (Denmark)
Bergami, Leonardo; Gaunaa, Mac
2010-01-01
This work presents a method to determine flutter and divergence instability limits for a two-dimensional (2-D) airfoil section fitted with an actively controlled trailing edge flap. This flap consists of a deformable trailing edge, which deformation is governed by control algorithms based...... on measurements of either heave displacement, local angle of attack or aerodynamic pressure difference measured over the airfoil. The purpose of the controlled deformable flap is to reduce fluctuations in the aerodynamic forces on the airfoil, which, according to recent studies, have a significant potential...... for fatigue load alleviation. The structural model of the 2-D airfoil section contains three degrees of freedom: heave translation, pitch rotation and flap deflection. A potential flow model provides the aerodynamic forces and their distribution. The unsteady aerodynamics are described using an indicial...
Extraction of airfoil data using PIV and pressure measurements
DEFF Research Database (Denmark)
Yang, Hua; Shen, Wen Zhong; Sørensen, Jens Nørkær
2011-01-01
A newly developed technique for determining the angle of attack (AOA) on a rotating blade is used to extract AOAs and airfoil data from measurements obtained during the MEXICO (Model rotor EXperiments in COntrolled conditions) rotor experiment. Detailed surface pressure and Particle Image...... airfoil data are compared to 2D data from wind tunnel experiments and XFOIL computations. The comparison suggests that the rotor is subject to severe 3D effects originating from the geometry of the rotor, and explains why the Blade Element Momentum technique with 2D airfoil data over‐predicts the loading...... is a reliable tool to extract airfoil data from experimental data. Copyright © 2010 John Wiley & Sons, Ltd....
Aerodynamic Analysis of Trailing Edge Enlarged Wind Turbine Airfoils
DEFF Research Database (Denmark)
Xu, Haoran; Shen, Wen Zhong; Zhu, Wei Jun
2014-01-01
characteristics of blunt trailing edge airfoils are caused by blunt body vortices at low angles of attack, and by the combined effect of separation and blunt body vortices at large angles of attack. With the increase of thickness of blunt trailing edge, the vibration amplitudes of lift and drag curves increase......The aerodynamic performance of blunt trailing edge airfoils generated from the DU- 91-W2-250, DU-97-W-300 and DU-96-W-350 airfoils by enlarging the thickness of trailing edge symmetrically from the location of maximum thickness to chord to the trailing edge were analyzed by using CFD and RFOIL...... methods at a chord Reynolds number of 3 × 106. The goal of this study is to analyze the aerodynamic performance of blunt trailing edge airfoils with different thicknesses of trailing edge and maximum thicknesses to chord. The steady results calculated by the fully turbulent k-ω SST, transitional k-ω SST...
Airfoil Pitch Control Using Trapped Vorticity Concentrations
Brzozowski, Daniel; Culp, John; Glezer, Ari
2007-11-01
Closed-loop feedback control of the attitude of a free pitching airfoil is effected without moving control surfaces by alternate actuation of nominally-symmetric trapped vorticity concentrations on the suction and pressure surfaces near the trailing edge. The pitching moment is varied with minimal lift and drag penalties over a broad range of angles of attack when the baseline flow is fully attached. Accumulation (trapping) and regulation of vorticity is managed by integrated hybrid actuators (each comprised of a miniature [O(0.01c)] obstruction and a synthetic jet actuator). In the present work, the model is trimmed using a position feedback loop and a servo motor actuator. Once the model is trimmed, the position feedback loop is opened and the servo motor acts like an inner loop control to alter the model's dynamic characteristics. Position control of the model is achieved using a reference model-based outer loop controller.
Symmetric airfoil geometry effects on leading edge noise.
Gill, James; Zhang, X; Joseph, P
2013-10-01
Computational aeroacoustic methods are applied to the modeling of noise due to interactions between gusts and the leading edge of real symmetric airfoils. Single frequency harmonic gusts are interacted with various airfoil geometries at zero angle of attack. The effects of airfoil thickness and leading edge radius on noise are investigated systematically and independently for the first time, at higher frequencies than previously used in computational methods. Increases in both leading edge radius and thickness are found to reduce the predicted noise. This noise reduction effect becomes greater with increasing frequency and Mach number. The dominant noise reduction mechanism for airfoils with real geometry is found to be related to the leading edge stagnation region. It is shown that accurate leading edge noise predictions can be made when assuming an inviscid meanflow, but that it is not valid to assume a uniform meanflow. Analytic flat plate predictions are found to over-predict the noise due to a NACA 0002 airfoil by up to 3 dB at high frequencies. The accuracy of analytic flat plate solutions can be expected to decrease with increasing airfoil thickness, leading edge radius, gust frequency, and Mach number.
Prediction of the wind turbine performance by using BEM with airfoil data extracted from CFD
DEFF Research Database (Denmark)
Yang, Hua; Shen, Wen Zhong; Xu, Haoran
2014-01-01
Blade element momentum (BEM) theory with airfoil data is a widely used technique for prediction of wind turbine aerodynamic performance, but the reliability of the airfoil data is an important factor for the prediction accuracy of aerodynamic loads and power. The airfoil characteristics used in BEM...... codes are mostly based on 2D wind tunnel measurements of airfoils with constant span. Due to 3D effects, a BEM code using airfoil data obtained directly from 2D wind tunnel measurements will not yield the correct loading and power. As a consequence, 2D airfoil characteristics have to be corrected before...
Hybrid inverse design method for nonlifting bodies in incompressible flow
CSIR Research Space (South Africa)
Broughton, BA
2006-11-01
Full Text Available A methodology for the inverse design of non-lifting axisymmetric bodies in compressible flow is presented. In this method, an inverse design approach based on conformal mapping is used to design a set of airfoils in isolation. These airfoils...
RANS Based Methodology for Predicting the Influence of Leading Edge Erosion on Airfoil Performance
Energy Technology Data Exchange (ETDEWEB)
Langel, Christopher M. [Univ. of California, Davis, CA (United States). Dept. of Mechanical and Aerospace Engineering; Chow, Raymond C. [Univ. of California, Davis, CA (United States). Dept. of Mechanical and Aerospace Engineering; van Dam, C. P. [Univ. of California, Davis, CA (United States). Dept. of Mechanical and Aerospace Engineering; Maniaci, David Charles [Sandia National Lab. (SNL-NM), Albuquerque, NM (United States). Wind Energy Technologies Dept.
2017-10-01
The impact of surface roughness on flows over aerodynamically designed surfaces is of interested in a number of different fields. It has long been known the surface roughness will likely accelerate the laminar- turbulent transition process by creating additional disturbances in the boundary layer. However, there are very few tools available to predict the effects surface roughness will have on boundary layer flow. There are numerous implications of the premature appearance of a turbulent boundary layer. Increases in local skin friction, boundary layer thickness, and turbulent mixing can impact global flow properties compounding the effects of surface roughness. With this motivation, an investigation into the effects of surface roughness on boundary layer transition has been conducted. The effort involved both an extensive experimental campaign, and the development of a high fidelity roughness model implemented in a R ANS solver. Vast a mounts of experimental data was generated at the Texas A&M Oran W. Nicks Low Speed Wind Tunnel for the calibration and validation of the roughness model described in this work, as well as future efforts. The present work focuses on the development of the computational model including a description of the calibration process. The primary methodology presented introduces a scalar field variable and associated transport equation that interacts with a correlation based transition model. The additional equation allows for non-local effects of surface roughness to be accounted for downstream of rough wall sections while maintaining a "local" formulation. The scalar field is determined through a boundary condition function that has been calibrated to flat plate cases with sand grain roughness. The model was initially tested on a NACA 0012 airfoil with roughness strips applied to the leading edge. Further calibration of the roughness model was performed using results from the companion experimental study on a NACA 63_{3} -418 airfoil
Geared-elevator flutter study. [transonic flutter characteristics of empennage
Ruhlin, C. L.; Doggett, R. V., Jr.; Gregory, R. A.
1976-01-01
The paper describes an experimental and analytical study of the transonic flutter characteristics of an empennage flutter model having an all-movable horizontal tail with a geared elevator. Two configurations were flutter tested: one with a geared elevator and one with a locked elevator with the model cantilever-mounted on a sting in the wind tunnel. The geared-elevator configuration fluttered experimentally at about 20% higher dynamic pressures than the locked-elevator configuration. The experimental flutter boundary was nearly flat at transonic speeds for both configurations. It was found that an analysis which treated the elevator as a discrete surface predicted flutter dynamic pressure levels better than analyses which treated the stabilizer and elevator as a warped surface. Warped-surface methods, however, predicted more closely the experimental flutter frequencies and Mach number trends.
Modern Schemes for Computation of Transonic Flows in Internal Aerodynamics
Directory of Open Access Journals (Sweden)
J. Fürst
2000-01-01
Full Text Available The work deals with numerical solution of steady transonic flows with applications in internal aerodynamics. The problems are described by the system of 2D or 3D Euler or Navier-Stokes equations. Our group developed during last years several central or upwind TVD finite volume methods for computation of transonic flows through a cascade or in a channel using grids of quadrilateral cells or triangular cells. We also developed other (non TVD schemes e.g. MacCormack scheme or several multistage Runge-Kutta finite volume schemes with applications of some acceleration techniques (multi-grid solution or hierarchical residual averaging. Our aim is to present some comparison of results of flows in internal aerodynamics using TVD schemes or other schemes. One can also compare the influence or artificial viscosity effects especially using TVD schemes.
Basic numerical methods. [of unsteady and transonic flow
Steger, Joseph L.; Van Dalsem, William R.
1989-01-01
Some of the basic finite-difference schemes that can be used to solve the nonlinear equations that describe unsteady inviscid and viscous transonic flow are reviewed. Numerical schemes for solving the unsteady Euler and Navier-Stokes, boundary-layer, and nonlinear potential equations are described. Emphasis is given to the elementary ideas used in constructing various numerical procedures, not specific details of any one procedure.
DETERMINATION OF PARASITE DRAG FOR COMMERCIAL AIRPLANE AT TRANSONIC SPEEDS
Directory of Open Access Journals (Sweden)
V. I. Shevyakov
2014-01-01
Full Text Available It considers the task of determination of parasite drag for subsonic commercial airplane on the base of theoretical calculation procedures. It contains the procedure for clarification of empirical approach to calculation of parasite drag at high speeds. It provides examples of calculation of parasite drag at transonic speeds as a result of step behind the slat on the wing of commercial airplane.
Transonic analysis of complex configurations using TRANAIR program
Saaris, G. R.; Gilkey, R. D.; Smit, K. L.; Tinoco, E. N.
1989-01-01
The application of a three-dimensional transonic flow analysis method, TRANAIR, is explored from the point of view of a user. Detailed features of the program are outlined to give a better understanding of capability. Numerous results are presented to show some of the complex configurations which have been analyzed. In particular, examples are provided which show the application to turbofan engine installation on transport aircraft.
On the general theory of thin airfoils for nonuniform motion
Reissner, Eric
1944-01-01
General thin-airfoil theory for a compressible fluid is formulated as boundary problem for the velocity potential, without recourse to the theory of vortex motion. On the basis of this formulation the integral equation of lifting-surface theory for an incompressible fluid is derived with the chordwise component of the fluid velocity at the airfoil as the function to be determined. It is shown how by integration by parts this integral equation can be transformed into the Biot-Savart theorem. A clarification is gained regarding the use of principal value definitions for the integral which occur. The integral equation of lifting-surface theory is used a s the starting point for the establishment of a theory for the nonstationary airfoil which is a generalization of lifting-line theory for the stationary airfoil and which might be called "lifting-strip" theory. Explicit expressions are given for section lift and section moment in terms of the circulation function, which for any given wing deflection is to be determined from an integral equation which is of the type of the equation of lifting-line theory. The results obtained are for airfoils of uniform chord. They can be extended to tapered airfoils. One of the main uses of the results should be that they furnish a practical means for the analysis of the aerodynamic span effect in the problem of wing flutter. The range of applicability of "lifting-strip" theory is the same as that of lifting-line theory so that its results may be applied to airfoils with aspect ratios as low as three.
Aeroelastic performance evaluation of a flexure box morphing airfoil concept
Pankonien, Alexander M.; Inman, Daniel J.
2014-04-01
The flexure-box morphing aileron concept utilizes Macro-Fiber Composites (MFCs) and a compliant box to create a conformal morphing aileron. This work evaluates the impact of the number of MFCs on the performance, power and mass of the aileron by experimentally investigating two different actuator configurations: unimorph and bimorph. Implemented in a NACA 0012 airfoil with 304.8 mm chord, the unimorph and bimorph configurations are experimentally tested over a range of flow speeds from 5 to 20 m/s and angles of attack from -20 to 20 degrees under aerodynamic loads in a wind tunnel. An embedded flexible sensor is installed in the aileron to evaluate the effect of aerodynamic loading on tip position. For both design choices, the effect of actuation on lift, drag and pitching moment coefficients are measured. Finally, the impact on aileron mass and average power consumption due to the added MFCs is considered. The results showed the unimorph exhibiting superior ability to influence flow up to 15 m/s, with equivalent power consumption and lower overall mass. At 20 m/s, the bimorph exhibited superior control over aerodynamic forces and the unimorph experienced significant deformation due to aerodynamic loading.
Singleton, Jeffrey D.; Yeager, William T., Jr.; Wilbur, Matthew L.
1990-01-01
An investigation was conducted in the NASA Langley Transonic Dynamics Tunnel to evaluate an advanced main rotor designed for use on a utility class helicopter, specifically the U.S. Army UH-60A Blackhawk. This rotor design incorporated advanced twist, airfoil cross sections, and geometric planform. For evaluation purposes, the current UH-60A main rotor was also tested and is referred to as the baseline blade set. A total of four blade sets were tested. One set of both the baseline and the advanced rotors were dynamically scaled to represent a full scale helicopter rotor blade design. The remaining advanced and baseline blade sets were not dynamically scaled so as to isolate the effects of structural elasticity. The investigation was conducted in hover and at rotor advance ratios ranging from 0.15 to 0.4 at a range of nominal test medium densities from 0.00238 to 0.009 slugs/cu ft. This range of densities, coupled with varying rotor lift and propulsive force, allowed for the simulation of several vehicle gross weight and density altitude combinations. Performance data are presented for all blade sets without analysis; however, cross referencing of data with flight condition may be useful to the analyst for validating aeroelastic theories and design methodologies as well as for evaluating advanced design parameters.
Aero-elastic stability of airfoil flow using 2-D CFD
Energy Technology Data Exchange (ETDEWEB)
Johansen, J. [Risoe National Lab., Roskilde (Denmark)
1999-03-01
A three degrees-of-freedom structural dynamics model has been coupled to a two-dimensional incompressible CFD code. The numerical investigation considers aero-elastic stability for two different airfoils; the NACA0012 and the LM 2 18 % airfoils. Stable and unstable configurations and limit cycle oscillations are predicted in accordance with literature for the first airfoil. An attempt to predict stall induced edge-wise vibrations on a wind turbine airfoil fails using this two-dimensional approach. (au)
Design and Optimization of Wings in Subsonic and Transonic Regime
2000-06-01
transici6n libre. In tinizaci6n. In 3 "r Congreso de M~todos Numie icos V Encontro Nacional de Mecanica Computacional , en Ingenierina, Zaragoza, Junio...on Computational Mechanics, Mecanica Computacional , Guimaraes, October,20- Buenos Aires, Argentina, June, 29-July 2,1998. 22,1997. 21-10 DISCUSSION
A computational procedure to improve airfoil performance considering shape and flow interactions
International Nuclear Information System (INIS)
Darbandi, M.; Taghvaey, M.J.; Schneider, G.E.
2004-01-01
Despite remarkable progress in shape design issue, there is still room to work on this topic considering different flow field conditions and specific aerodynamic applications. Today, the optimization techniques are known as a major tool to reach the best possible aerodynamic shape for some specific conditions. In general direct optimization techniques, the optimization process is started from choosing a suitable primitive shape and the shape is improved by suitable considerations of the design objectives and constraints. In a similar attempt, we develop a new optimization strategy to improve the airfoil shape for specified applications. The strategy involves several stages. It includes to determine the flow conditions and design parameters, to establish the objective function, to select a suitable primitive shape, to generate a mechanism for inserting gradual shape changes, to generate grids around each defined shape, to solve the flow field for each separate shape, to collect the solution data, to change the discrete data to the continuous distribution functions, to construct the objective function, and to minimize the objective function using the steepest descent approach. No constraint function is incorporated into the constructed objective function. The cruise flight of an aircraft at an specified altitude is supposed to be the flow field conditions around the proposed airfoil. Nevertheless, the flow field is assumed to be viscous and compressible as well as turbulent. The procedure is tested starting from two generic airfoil shapes with and without camber. The developed algorithm works well for both cases; however, it may not converge to identical shapes if the primitive shapes are not identical. (author)
Computational Analysis of Dual Radius Circulation Control Airfoils
Lee-Rausch, E. M.; Vatsa, V. N.; Rumsey, C. L.
2006-01-01
The goal of the work is to use multiple codes and multiple configurations to provide an assessment of the capability of RANS solvers to predict circulation control dual radius airfoil performance and also to identify key issues associated with the computational predictions of these configurations that can result in discrepancies in the predicted solutions. Solutions were obtained for the Georgia Tech Research Institute (GTRI) dual radius circulation control airfoil and the General Aviation Circulation Control (GACC) dual radius airfoil. For the GTRI-DR airfoil, two-dimensional structured and unstructured grid computations predicted the experimental trend in sectional lift variation with blowing coefficient very well. Good code to code comparisons between the chordwise surface pressure coefficients and the solution streamtraces also indicated that the detailed flow characteristics were matched between the computations. For the GACC-DR airfoil, two-dimensional structured and unstructured grid computations predicted the sectional lift and chordwise pressure distributions accurately at the no blowing condition. However at a moderate blowing coefficient, although the code to code variation was small, the differences between the computations and experiment were significant. Computations were made to investigate the sensitivity of the sectional lift and pressure distributions to some of the experimental and computational parameters, but none of these could entirely account for the differences in the experimental and computational results. Thus, CFD may indeed be adequate as a prediction tool for dual radius CC flows, but limited and difficult to obtain two-dimensional experimental data prevents a confident assessment at this time.
Active Control of Aerodynamic Forces on a Rapidly Maneuvering Airfoil
Brzozowski, Daniel; Culp, John; Glezer, Ari
2009-11-01
The unsteady aerodynamic forces and moments on a rapidly maneuvering free-moving airfoil are investigated in wind tunnel experiments. The airfoil is mounted on a 2-DOF traverse and its trim and dynamic characteristics are controlled using position and attitude feedback loops that are actuated by servo motors. The motion of the airfoil is effected by bi-directional changes in the pitching moment using controllable trapped vorticity concentrations on both the suction and pressure surfaces near the trailing edge that are induced and regulated by hybrid synthetic jet actuators. The dynamic coupling between the actuation and the time-dependent flow field is characterized using simultaneous force and velocity measurements that are taken phase-locked to the commanded actuation waveform. The unsteady flow characteristics induced by the fluidic actuation during a prescribed maneuver are compared with the effects of a simple rigid-body motion of the airfoil when an external torque is used to achieve a similar maneuver. It is shown that the time-dependent aerodynamic forces and induced flow fields in the near wake of the moving airfoil are significantly different, emphasizing the role of the coupling between the flow control actuation and the model's unsteady aerodynamics.
Miotto, Renato Fuzaro; Wolf, William Roberto; de Santana, Leandro Dantas
2016-01-01
Current knowledge on the noise generation mechanisms of an airfoil subjected to a turbulent flow indicates that an increment to the airfoil thickness leads to a reduction of the leading-edge noise. This effect is generally attributed to the turbulence distortion occurring close upstream the airfoil
Aerodynamic characteristics of wind turbine blade airfoils at high angles-of-attack
Timmer, W.A.
2010-01-01
Airfoil characteristics at deep stall angles were investigated. It appeared that the maximum drag coefficient as a function of the airfoil upwind y/c ordinate at x/c=0.0125 can be approximated by a straight line. The lift-drag ratios in deep stall of a number of airfoils with moderate lower surface
Improved Swimming Performance in Hydrodynamically- coupled Airfoils
Heydari, Sina; Shelley, Michael J.; Kanso, Eva
2017-11-01
Collective motion is a widespread phenomenon in the animal kingdom from fish schools to bird flocks. Half of the known fish species are thought to exhibit schooling behavior during some phase of their life cycle. Schooling likely occurs to serve multiple purposes, including foraging for resources and protection from predators. Growing experimental and theoretical evidence supports the hypothesis that fish can benefit from the hydrodynamic interactions with their neighbors, but it is unclear whether this requires particular configurations or regulations. Here, we propose a physics-based approach that account for hydrodynamic interactions among swimmers based on the vortex sheet model. The benefit of this model is that it is scalable to a large number of swimmers. We start by examining the case of two swimmers, heaving plates, moving in parallel and in tandem. We find that for the same heaving amplitude and frequency, the coupled-swimmers move faster and more efficiently. This increase in velocity depends strongly on the configuration and separation distance between the swimmers. Our results are consistent with recent experimental findings on heaving airfoils and underline the role of fluid dynamic interactions in the collective behavior of swimmers.
Uncertainty Analysis for a Jet Flap Airfoil
Green, Lawrence L.; Cruz, Josue
2006-01-01
An analysis of variance (ANOVA) study was performed to quantify the potential uncertainties of lift and pitching moment coefficient calculations from a computational fluid dynamics code, relative to an experiment, for a jet flap airfoil configuration. Uncertainties due to a number of factors including grid density, angle of attack and jet flap blowing coefficient were examined. The ANOVA software produced a numerical model of the input coefficient data, as functions of the selected factors, to a user-specified order (linear, 2-factor interference, quadratic, or cubic). Residuals between the model and actual data were also produced at each of the input conditions, and uncertainty confidence intervals (in the form of Least Significant Differences or LSD) for experimental, computational, and combined experimental / computational data sets were computed. The LSD bars indicate the smallest resolvable differences in the functional values (lift or pitching moment coefficient) attributable solely to changes in independent variable, given just the input data points from selected data sets. The software also provided a collection of diagnostics which evaluate the suitability of the input data set for use within the ANOVA process, and which examine the behavior of the resultant data, possibly suggesting transformations which should be applied to the data to reduce the LSD. The results illustrate some of the key features of, and results from, the uncertainty analysis studies, including the use of both numerical (continuous) and categorical (discrete) factors, the effects of the number and range of the input data points, and the effects of the number of factors considered simultaneously.
Directory of Open Access Journals (Sweden)
YOUNES EL KHCHINE
2017-11-01
Full Text Available The design of rotor blades has a great effect on the aerodynamics performances of horizontal axis wind turbine and its efficiency. This work presents the effects of mesh refinement and boundary layer on aerodynamic performances of wind turbine S809 rotor. Furthermore, the simulation of fluid flow is taken for S809 airfoil wind turbine blade using ANSYS/FLUENT software. The problem is solved by the conservation of mass and momentum equations for unsteady and incompressible flow using advanced SST k-ω turbulence model, in order to predict the effects of mesh refinement and boundary layer on aerodynamics performances. Lift and drag coefficients are the most important parameters in studying the wind turbine performance, these coefficients are calculated for four meshes refinement and different angles of attacks with Reynolds number is 106. The study is applied to S809 airfoil which has 21% thickness, specially designed by NREL for horizontal axis wind turbines.
Pressure distribution over an NACA 23012 airfoil with a fixed slot and a slotted flap
Harris, Thomas A; Lowry, John G
1942-01-01
Report presents the results of a pressure-distribution investigation conducted in the Langley Memorial Aeronautical Laboratory 7 by 10-foot wind tunnel to determine the air loads on an NACA 23012 airfoil in combination with a fixed leading-edge slot and a slotted flap. Pressures were measured over the upper and lower surfaces of the component parts of the combination for several angles of attack and at several flap settings. The data, presented as pressure diagrams and graphs of section coefficients, are applicable to rib, slat, and flap designs for the combination.
Laser anemometer measurements in a transonic axial-flow fan rotor
Strazisar, Anthony J.; Wood, Jerry R.; Hathaway, Michael D.; Suder, Kenneth L.
1989-01-01
Laser anemometer surveys were made of the 3-D flow field in NASA rotor 67, a low aspect ratio transonic axial-flow fan rotor. The test rotor has a tip relative Mach number of 1.38. The flowfield was surveyed at design speed at near peak efficiency and near stall operating conditions. Data is presented in the form of relative Mach number and relative flow angle distributions on surfaces of revolution at nine spanwise locations evenly spaced from hub to tip. At each spanwise location, data was acquired upstream, within, and downstream of the rotor. Aerodynamic performance measurements and detailed rotor blade and annulus geometry are also presented so that the experimental results can be used as a test case for 3-D turbomachinery flow analysis codes.
Nonlinear dynamics approach of modeling the bifurcation for aircraft wing flutter in transonic speed
DEFF Research Database (Denmark)
Matsushita, Hiroshi; Miyata, T.; Christiansen, Lasse Engbo
2002-01-01
The procedure of obtaining the two-degrees-of-freedom, finite dimensional. nonlinear mathematical model. which models the nonlinear features of aircraft flutter in transonic speed is reported. The model enables to explain every feature of the transonic flutter data of the wind tunnel tests...
2015-10-01
Aerodynamic Experimental Facility (AEF) by Charith R Ranawake Approved for public release; distribution is unlimited...Laboratory The Automation of the Transonic Experimental Facility (TEF) and the Aerodynamic Experimental Facility (AEF) by Charith R Ranawake Weapons...To) 05/2015–08/2015 4. TITLE AND SUBTITLE The Automation of the Transonic Experimental Facility (TEF) and the Aerodynamic Experimental Facility
Investigation of Airfoil Aeroacoustics with Blowing Control Mechanism
Directory of Open Access Journals (Sweden)
Baha ZAFER
2016-11-01
Full Text Available In this investigation, it is dealt with computational aero-acoustic analysis of an airfoil with jet blowing. The airfoil shape is selected as NACA0015 profile with jet blowing on upper surface. The calculations of analysis are done by using commercial finite volume solver. The k-ε turbulence model is used for the turbulence modeling and the Ffowcs Williams and Hawking acoustic analogy model is run for determination of acoustic data. The numerical results are compared with experimental data for computed Sound Pressure Level without jet blowing and well agreement is observed. In the case of jet blowing, the effects of different jet angle, velocity ratio and angle of attack on airfoil are investigated and noise levels of non jet cases and jet blowing cases are studied.
Optimization of Flapping Airfoils for Maximum Thrust and Propulsive Efficiency
Directory of Open Access Journals (Sweden)
I. H. Tuncer
2004-01-01
Full Text Available A numerical optimization algorithm based on the steepest decent along the variation of the optimization function is implemented for maximizing the thrust and/or propulsive efficiency of a single flapping airfoil. Unsteady, low speed laminar flows are computed using a Navier-Stokes solver on moving overset grids. The flapping motion of the airfoil is described by a combined sinusoidal plunge and pitching motion. Optimization parameters are taken to be the amplitudes of the plunge and pitching motions, and the phase shift between them. Computations are performed in parallel in a work station cluster. The numerical simulations show that high thrust values may be obtained at the expense of reduced efficiency. For high efficiency in thrust generation, the induced angle of attack of the airfoil is reduced and large scale vortex formations at the leading edge are prevented.
Response of a thin airfoil encountering strong density discontinuity
Energy Technology Data Exchange (ETDEWEB)
Marble, F.E. [California Inst. of Tech., Pasadena, CA (United States). Karman Lab. of Fluid Mechanics
1993-12-01
Airfoil theory for unsteady motion has been developed extensively assuming the undisturbed medium to be of uniform density, a restriction accurate for motion in the atmosphere. In some instances, notably for airfoil comprising fan, compressor and turbine blade rows, the undisturbed medium may carry density variations or ``spots``, resulting from non-uniformities in temperature or composition, of a size comparable to the blade chord. This condition exists for turbine blades, immediately downstream of the main burner of a gas turbine engine where the density fluctuations of the order of 50 percent may occur. Disturbances of a somewhat smaller magnitude arise from the ingestion of hot boundary layers into fans, and exhaust into hovercraft. Because these regions of non-uniform density convect with the moving medium, the airfoil experiences a time varying load and moment which the authors calculate.
1990-11-01
des pressions ct des vitesses. en tenant compte des performances souhait~es et du processus des pertes au nivcao de la couche timite. Loptimisation...directe qoi r~sutte du calcul inverse de la couche timite est d&crite, ainsi que Coptimisation itirative obtenue par minimalisation des pertes . Les deux...addition, the several cross numerical tools to be used in the design procedure. 4, .1-14 10. REFERENCES 13. Polito, L., "Un Metodo Esatto per it
Trailing edge noise model applied to wind turbine airfoils
Energy Technology Data Exchange (ETDEWEB)
Bertagnolio, F.
2008-01-15
The aim of this work is firstly to provide a quick introduction to the theory of noise generation that are relevant to wind turbine technology with focus on trailing edge noise. Secondly, the socalled TNO trailing edge noise model developed by Parchen [1] is described in more details. The model is tested and validated by comparing with other results from the literature. Finally, this model is used in the optimization process of two reference airfoils in order to reduce their noise signature: the RISOE-B1-18 and the S809 airfoils. (au)
Unsteady Surface Pressure Measurements on a Pitching Airfoil
1985-03-12
through 8 Dynamics 7512B amplifiers. The pitching motions of the airfoil were generated by 6°jN\\! 920O/_ a PDP 11/03 computer controlling a Control...acquisition system. The pressure data were used to calculate pressure 2 coefficients which were in turn integrated to compute lift coefficients. Both...Airfoils," AIAA J., Vol. 13, No. 1, 17. Gormont, R.E., "A Mathenatical Model pp 71-79, Jan 1975. of Unsteady Aerodynamics and Radial 4. McAlister, K.W
A dynamic stall model for airfoils with deformable trailing edges
DEFF Research Database (Denmark)
Andersen, Peter Bjørn; Gaunaa, Mac; Bak, Dan Christian
2007-01-01
on an airfoil section undergoing arbitrary motion in heave, lead-lag, pitch, Trailing Edge (TE) flapping. In the linear region, the model reduces to the inviscid model of Gaunaa [4], which includes the aerodynamic effect of a thin airfoil with a deformable camberline in inviscid flow. Therefore, the proposed......The present work contains an extension of the Beddoes-Leishman (B-L) type dynamic stall model, as described by Hansen et al. [7]. In this work a Deformable Trailing Edge Geometry (DTEG) has been added to the dynamic stall model. The model predicts the unsteady aerodynamic forces and moments...
NUMERICAL INVESTIGATION OF TWO ELEMENT CAMBER MORPHING AIRFOIL IN LOW REYNOLDS NUMBER FLOWS
Directory of Open Access Journals (Sweden)
RAJESH SENTHIL KUMAR T.
2017-07-01
Full Text Available Aerodynamic performance of a two-element camber morphing airfoil was investigated at low Reynolds number using the transient SST model in ANSYS FLUENT 14.0 and eN method in XFLR5. The two-element camber morphing concept was employed to morph the baseline airfoil into another airfoil by altering the orientation of mean-line at 35% of the chord to achieve better aerodynamic efficiency. NACA 0012 was selected as baseline airfoil. NACA 23012 was chosen as the test case as it has the camber-line similar to that of the morphed airfoil and as it has the same thickness as that of the baseline airfoil. The simulations were carried out at chord based Reynolds numbers of 2.5×105 and 3.9×105. The aerodynamic force coefficients, aerodynamic efficiency and the location of the transition point of laminar separation bubble over these airfoils were studied for various angles of attack. It was found that the aerodynamic efficiency of the morphed airfoil was 12% higher than that of the target airfoil at 4° angle of attack for Reynolds number of 3.9×105 and 54% rise in aerodynamic performance was noted as Reynolds number was varied from 2.5×105 to 3.9×105. The morphed airfoil exhibited the nature of low Reynolds number airfoil.
Study of Near-Stall Flow Behavior in a Modern Transonic Fan with Composite Sweep
Hah, Chunill; Shin, Hyoun-Woo
2011-01-01
Detailed flow behavior in a modern transonic fan with a composite sweep is investigated in this paper. Both unsteady Reynolds-averaged Navier-Stokes (URANS) and Large Eddy Simulation (LES) methods are applied to investigate the flow field over a wide operating range. The calculated flow fields are compared with the data from an array of high-frequency response pressure transducers embedded in the fan casing. The current study shows that a relatively fine computational grid is required to resolve the flow field adequately and to calculate the pressure rise across the fan correctly. The calculated flow field shows detailed flow structure near the fan rotor tip region. Due to the introduction of composite sweep toward the rotor tip, the flow structure at the rotor tip is much more stable compared to that of the conventional blade design. The passage shock stays very close to the leading edge at the rotor tip even at the throttle limit. On the other hand, the passage shock becomes stronger and detaches earlier from the blade passage at the radius where the blade sweep is in the opposite direction. The interaction between the tip clearance vortex and the passage shock becomes intense as the fan operates toward the stall limit, and tip clearance vortex breakdown occurs at near-stall operation. URANS calculates the time-averaged flow field fairly well. Details of measured RMS static pressure are not calculated with sufficient accuracy with URANS. On the other hand, LES calculates details of the measured unsteady flow features in the current transonic fan with composite sweep fairly well and reveals the flow mechanism behind the measured unsteady flow field.
HSCT Ref-H Transonic Flap Data Base: Wind-Tunnel Test and Comparison with Theory
Vijgen, Paul M.
1999-01-01
In cooperation with personnel from the Boeing ANP Laboratory and NASA Langley, a performance test was conducted using the Reference-H 1.675% model ("NASA Modular Model") without nacelles at the NASA Langley 16-Ft Transonic Tunnel. The main objective of the test was to determine the drag reduction achievable with leading-edge and trailing-edge flaps deflected along the outboard wing span at transonic Mach numbers (M = 0.9 to 1.2) for purpose of preliminary design and for comparison with computational predictions. The obtained drag data with flap deflections for Mach numbers of 1.07 to 1.20 are unique for the Reference H wing. Four leading-edge and two trailing-edge flap deflection angles were tested at a mean-wing chord-Reynolds number of about 5.7 million. An outboard-wing leading-edge flap deflection of 81 provides a 4.5 percent drag reduction at M = 1.2 A = 0.2), and much larger values at lower Mach numbers with larger flap deflections. The present results for the baseline (no flaps deflected) compare reasonably well with previous Boeing and NASA Ref-H tunnel tests, including high-Reynolds number NTF results. Viscous CFD simulations using the OVERFLOW thin-layer N.S. method properly predict the observed trend in drag reduction at M = 1.2 as function of leading-edge flap deflection. Modified linear theory properly predicts the flap effects on drag at subsonic conditions (Aero2S code), and properly predicts the absolute drag for the 40 and 80 leading-edge deflection at M = 1.2 (A389 code).
Rotational Augmentation on a 2.3 MW Rotor Blade with Thick Flatback Airfoil Cross-Sections: Preprint
Energy Technology Data Exchange (ETDEWEB)
Schreck, S.; Fingersh, L.; Siegel, K.; Singh, M.; Medina, P.
2013-01-01
Rotational augmentation was analyzed for a 2.3 MW wind turbine, which was equipped with thick flatback airfoils at inboard radial locations and extensively instrumented for acquisition of time varying surface pressures. Mean aerodynamic force and surface pressure data were extracted from an extensive field test database, subject to stringent criteria for wind inflow and turbine operating conditions. Analyses of these data showed pronounced amplification of aerodynamic forces and significant enhancements to surface pressures in response to rotational influences, relative to two-dimensional, stationary conditions. Rotational augmentation occurrence and intensity in the current effort was found to be consistent with that observed in previous research. Notably, elevated airfoil thickness and flatback design did not impede rotational augmentation.
Aerodynamic behavior of an airfoil with morphing trailing edge for wind turbine applications
International Nuclear Information System (INIS)
Wolff, T; Ernst, B; Seume, J R
2014-01-01
The length of wind turbine rotor blades has been increased during the last decades. Higher stresses arise especially at the blade root because of the longer lever arm. One way to reduce unsteady blade-root stresses caused by turbulence, gusts, or wind shear is to actively control the lift in the blade tip region. One promising method involves airfoils with morphing trailing edges to control the lift and consequently the loads acting on the blade. In the present study, the steady and unsteady behavior of an airfoil with a morphing trailing edge is investigated. Two-dimensional Reynolds-Averaged Navier-Stokes (RANS) simulations are performed for a typical thin wind turbine airfoil with a morphing trailing edge. Steady-state simulations are used to design optimal geometry, size, and deflection angles of the morphing trailing edge. The resulting steady aerodynamic coefficients are then analyzed at different angles of attack in order to determine the effectiveness of the morphing trailing edge. In order to investigate the unsteady aerodynamic behavior of the optimal morphing trailing edge, time- resolved RANS-simulations are performed using a deformable grid. In order to analyze the phase shift between the variable trailing edge deflection and the dynamic lift coefficient, the trailing edge is deflected at four different reduced frequencies for each different angle of attack. As expected, a phase shift between the deflection and the lift occurs. While deflecting the trailing edge at angles of attack near stall, additionally an overshoot above and beyond the steady lift coefficient is observed and evaluated
Aerodynamic behavior of an airfoil with morphing trailing edge for wind turbine applications
Wolff, T.; Ernst, B.; Seume, J. R.
2014-06-01
The length of wind turbine rotor blades has been increased during the last decades. Higher stresses arise especially at the blade root because of the longer lever arm. One way to reduce unsteady blade-root stresses caused by turbulence, gusts, or wind shear is to actively control the lift in the blade tip region. One promising method involves airfoils with morphing trailing edges to control the lift and consequently the loads acting on the blade. In the present study, the steady and unsteady behavior of an airfoil with a morphing trailing edge is investigated. Two-dimensional Reynolds-Averaged Navier-Stokes (RANS) simulations are performed for a typical thin wind turbine airfoil with a morphing trailing edge. Steady-state simulations are used to design optimal geometry, size, and deflection angles of the morphing trailing edge. The resulting steady aerodynamic coefficients are then analyzed at different angles of attack in order to determine the effectiveness of the morphing trailing edge. In order to investigate the unsteady aerodynamic behavior of the optimal morphing trailing edge, time- resolved RANS-simulations are performed using a deformable grid. In order to analyze the phase shift between the variable trailing edge deflection and the dynamic lift coefficient, the trailing edge is deflected at four different reduced frequencies for each different angle of attack. As expected, a phase shift between the deflection and the lift occurs. While deflecting the trailing edge at angles of attack near stall, additionally an overshoot above and beyond the steady lift coefficient is observed and evaluated.
Effects of grit roughness and pitch oscillations on the S814 airfoil
Energy Technology Data Exchange (ETDEWEB)
Janiszewska, J.M.; Ramsay, R.R.; Hoffmann, M.J.; Gregorek, G.M. [Ohio State Univ., Columbus, OH (United States)
1996-07-01
Horizontal-axis wind turbine rotors experience unsteady aerodynamics when the rotor is yawed, when rotor blades pass through the support tower wake, and when the wind is gusting. An understanding of this unsteady behavior is necessary to assist in the design of new rotor airfoils. The rotors also experience performance degradation due to surface roughness. These surface irregularities are due to the accumulation of insect debris, ice, and/or the aging process. Wind tunnel studies that examine both the steady and unsteady behavior of airfoils can help define pertinent flow phenomena, and the resultant data can also be used to validate analytical computer codes. An S814 airfoil model was tested in The Ohio State University Aeronautical and Astronautical Research Laboratory (OSU/AARL) 3 X 5 subsonic wind tunnel (3 X 5) under steady flow with both stationary model conditions and pitch oscillations. To study the extent of performance loss due to surface roughness, a leading edge grit roughness pattern (LEGR) was used to simulate leading edge contamination. After baseline cases were completed, the LEGR was applied for both steady state and model pitch oscillation cases. The Reynolds numbers for steady state conditions were 0.75, 1, 1.25 and 1.5 million, while the angle of attack ranged from -20{degrees} to +40{degrees}. While the model underwent pitch oscillations, data were acquired at Reynolds numbers of 0.75, 1, 1.25, and 1.5 million, at frequencies of 0.6, 1.2, and 1.8 Hz. Two sine wave forcing functions {+-}5.5{degrees} and {+-}10{degrees}, were used; at mean angles of attack of 8{degrees}, 14{degrees}, and 20{degrees}. For purposes herein, any reference to unsteady conditions means the model was in pitch oscillation.
Aeroelastic Tailoring of Transport Wings Including Transonic Flutter Constraints
Stanford, Bret K.; Wieseman, Carol D.; Jutte, Christine V.
2015-01-01
Several minimum-mass optimization problems are solved to evaluate the effectiveness of a variety of novel tailoring schemes for subsonic transport wings. Aeroelastic stress and panel buckling constraints are imposed across several trimmed static maneuver loads, in addition to a transonic flutter margin constraint, captured with aerodynamic influence coefficient-based tools. Tailoring with metallic thickness variations, functionally graded materials, balanced or unbalanced composite laminates, curvilinear tow steering, and distributed trailing edge control effectors are all found to provide reductions in structural wing mass with varying degrees of success. The question as to whether this wing mass reduction will offset the increased manufacturing cost is left unresolved for each case.
Experimental and computational study of transonic flow about swept wings
Bertelrud, A.; Bergmann, M. Y.; Coakley, T. J.
1980-01-01
An experimental investigation of NACA 0010 and 10% circular arc wing models, swept at 45 deg, spanning a channel, and at zero angle of attack is described. Measurements include chordwise and spanwise surface pressure distributions and oil-flow patterns for a range of transonic Mach numbers and Reynolds numbers. Calculations using a new three-dimensional Navier-Stokes code and a two-equation turbulence model are included for the circular-arc wing flow. Reasonable agreement between measurements and computations is obtained.
Special Course on Subsonic/Transonic Aerodynamic Interference for Aircraft
1983-07-01
1982 11. H. Jakob Erweiterung eines 2-D Panelverfahrens auf Profile mit dicken Hinterkanten und Kopplung des Verfahrens mit einem...number of years. Indeed there are NACA reports by Nielsen at al written in the mid to late 50’s which used the basic approach still given in DATC0M...K.D.Klevenhusen Calculation of wing-body-nacelle interference in subsonic and transonic potential H. Jakob flow. AGARD-CP-301, May 1981. H. Struck 12 J. Barche
Transonic twins in 3D bcc iron crystal
Czech Academy of Sciences Publication Activity Database
Spielmannová, Alena; Machová, Anna; Hora, Petr
2010-01-01
Roč. 48, č. 2 (2010), s. 296-302 ISSN 0927-0256 R&D Projects: GA AV ČR KJB200760802; GA ČR GA101/09/1630 Institutional research plan: CEZ:AV0Z20760514 Keywords : transonic twins * bcc iron * molecular dynamic simulation Subject RIV: JG - Metallurgy Impact factor: 1.458, year: 2010 http://apps.isiknowledge.com/full_record.do?product=WOS&search_mode=GeneralSearch&qid=3&SID=V1mj77dMKmjeKefm7Db&page=1&doc=1
Detached Eddy Simulations of an Airfoil in Turbulent Inflow
DEFF Research Database (Denmark)
Gilling, Lasse; Sørensen, Niels; Davidson, Lars
2009-01-01
The effect of resolving inflow turbulence in detached eddy simulations of airfoil flows is studied. Synthetic turbulence is used for inflow boundary condition. The generated turbulence fields are shown to decay according to experimental data as they are convected through the domain with the free ...
CFD code comparison for 2D airfoil flows
DEFF Research Database (Denmark)
Sørensen, Niels N.; Méndez, B.; Muñoz, A.
2016-01-01
The current paper presents the effort, in the EU AVATAR project, to establish the necessary requirements to obtain consistent lift over drag ratios among seven CFD codes. The flow around a 2D airfoil case is studied, for both transitional and fully turbulent conditions at Reynolds numbers of 3 × ...
Overview of results from 2D airfoil testing at Risoe
Energy Technology Data Exchange (ETDEWEB)
Fuglsang, P. [Risoe National Lab., Wind Energy and Atmospheric Physics Dept., Roskilde (Denmark)
1997-12-31
This paper gives an overview of the results from two dimensional airfoil testing at Risoe. A two dimensional testing method was recently developed where a test rig is inserted into an open jet flow in a wind tunnel of the close return loop type with an open test section. Pressure measurements provide the lift and drag forces. Both stationary flow and dynamic inflow from pitch motion are possible. The wind tunnel static pressure and total dynamic pressures were calibrated and wind tunnel boundary corrections were found. So far, the testing method was verified by comparison of NACA 63-215 airfoil measurements to numerical predictions and to measurements. Furthermore, the Risoe-1, FFA-W3-241, FFA-W3-301 and NACA 63-430 airfoils were measured. Different types of leading edge roughness and vortex generators were investigated. For all airfoils, good agreements with predictions were obtained on both pressure distribution and on lift coefficient. The drag coefficients were slightly higher than predicted. (eg) 10 refs.
Numerical simulation of airfoil trailing edge serration noise
DEFF Research Database (Denmark)
Zhu, Wei Jun; Shen, Wen Zhong
In the present work, numerical simulations are carried out for a low noise airfoil with and without serrated Trailing Edge. The Ffowcs Williams-Hawkings acoustic analogy is implemented into the in-house incompressible flow solver EllipSys3D. The instantaneous hydrodynamic pressure and velocity...
Flow characteristics over NACA4412 airfoil at low Reynolds number
Directory of Open Access Journals (Sweden)
Genç Mustafa Serdar
2016-01-01
Full Text Available In this study, the flow phenomena over NACA4412 were experimentally observed at various angle of attack and Reynolds number of 25000, 50000 and 75000, respectively. NACA4412 airfoil was manufactured at 3D printer and each tips of the wing were closed by using plexiglas to obtain two-dimensional airfoil. The experiments were conducted at low speed wind tunnel. The force measurement and hot-wire experiments were conducted to obtain data so that the flow phenomenon at the both top and bottom of the airfoil such as the flow separation and vortex shedding were observed. Also, smoke-wire experiment was carried out to visualize the surface flow pattern. After obtaining graphics from both force measurement experiment and hot-wire experiment compared with smoke wire experiment, it was noticed that there is a good coherence among the experiments. It was concluded that as Re number increased, the stall angle increased. And the separation bubble moved towards leading edge over the airfoil as the angle of attack increased.
Unsteady Double Wake Model for the Simulation of Stalled Airfoils
DEFF Research Database (Denmark)
Ramos García, Néstor; Cayron, Antoine; Sørensen, Jens Nørkær
2015-01-01
separation and its dynamics. In this paper, the calculated integral forces have been successfully validated against wind tunnel measurements for the FFA-W3-211 airfoil. Furthermore, the computed highly unsteady flow field is analyzed in detail for a set of angles of attack ranging from light to deep stall...
Mechanism of unconventional aerodynamic characteristics of an elliptic airfoil
Directory of Open Access Journals (Sweden)
Sun Wei
2015-06-01
Full Text Available The aerodynamic characteristics of elliptic airfoil are quite different from the case of conventional airfoil for Reynolds number varying from about 104 to 106. In order to reveal the fundamental mechanism, the unsteady flow around a stationary two-dimensional elliptic airfoil with 16% relative thickness has been simulated using unsteady Reynolds-averaged Navier–Stokes equations and the γ-Reθt‾ transition turbulence model at different angles of attack for flow Reynolds number of 5 × 105. The aerodynamic coefficients and the pressure distribution obtained by computation are in good agreement with experimental data, which indicates that the numerical method works well. Through this study, the mechanism of the unconventional aerodynamic characteristics of airfoil is analyzed and discussed based on the computational predictions coupled with the wind tunnel results. It is considered that the boundary layer transition at the leading edge and the unsteady flow separation vortices at the trailing edge are the causes of the case. Furthermore, a valuable insight into the physics of how the flow behavior affects the elliptic airfoil’s aerodynamics is provided.
Advanced turbine study. [airfoil coling in rocket turbines
1982-01-01
Experiments to determine the available increase in turbine horsepower achieved by increasing turbine inlet temperature over a range of 1800 to 2600 R, while applying current gas turbine airfoil cling technology are discussed. Four cases of rocket turbine operating conditions were investigated. Two of the cases used O2/H2 propellant, one with a fuel flowrate of 160 pps, the other 80 pps. Two cases used O2/CH4 propellant, each having different fuel flowrates, pressure ratios, and inlet pressures. Film cooling was found to be the required scheme for these rocket turbine applications because of the high heat flux environments. Conventional convective or impingement cooling, used in jet engines, is inadequate in a rocket turbine environment because of the resulting high temperature gradients in the airfoil wall, causing high strains and low cyclic life. The hydrogen-rich turbine environment experienced a loss, or no gain, in delivered horsepower as turbine inlet temperature was increased at constant airfoil life. The effects of film cooling with regard to reduced flow available for turbine work, dilution of mainstream gas temperature and cooling reentry losses, offset the relatively low specific work capability of hydrogen when increasing turbine inlet temperature over the 1800 to 2600 R range. However, the methane-rich environment experienced an increase in delivered horsepower as turbine inlet temperature was increased at constant airfoil life. The results of a materials survey and heat transfer and durability analysis are discussed.
Halle, J. E.; Ruschak, J. T.
1975-01-01
A highly loaded, high tip-speed fan rotor was designed with multiple-circular-arc airfoil sections as a replacement for a marginally successful rotor which had precompression airfoil sections. The substitution of airfoil sections was the only aerodynamic change. Structural design of the redesigned rotor blade was guided by successful experience with the original blade. Calculated stress levels and stability parameters for the redesigned rotor are within limits demonstrated in tests of the original rotor.
Energy Technology Data Exchange (ETDEWEB)
Suder, K.L. [NASA-Lewis Research Center, Cleveland, OH (United States); Celestina, M.L. [Sverdrup Technology Inc., Cleveland, OH (United States). Lewis Research Center Group
1996-04-01
Experimental and computational techniques are used to investigate tip clearance flows in a transonic axial compressor rotor at design and part-speed conditions. Laser anemometer data acquired in the endwall region are presented for operating conditions near peak efficiency and near stall at 100 percent design speed and at near peak efficiency at 60 percent design speed. The role of the passage shock/leakage vortex interaction at design speed, the radial influence of the tip clearance flow extends to 20 times the physical tip clearance height. At part speed, in the absence of the shock, the radial extent is only five times the tip clearance height. Both measurements and analysis indicate that under part-speed operating conditions a second vortex, which does not originate from the tip leakage flow, forms in the end-wall region within the blade passage and exits the passage near midpitch. Mixing of the leakage vortex with the primary flow downstream of the rotor at both design and part-speed conditions is also discussed.
Suder, Kenneth L.; Celestina, Mark L.
1995-01-01
Experimental and computational techniques are used to investigate tip clearance flows in a transonic axial compressor rotor at design and part speed conditions. Laser anemometer data acquired in the endwall region are presented for operating conditions near peak efficiency and near stall at 100% design speed and at near peak efficiency at 60% design speed. The role of the passage shock/leakage vortex interaction in generating endwall blockage is discussed. As a result of the shock/vortex interaction at design speed, the radial influence of the tip clearance flow extends to 20 times the physical tip clearance height. At part speed, in the absence of the shock, the radial extent is only 5 times the tip clearance height. Both measurements and analysis indicate that under part-speed operating conditions a second vortex, which does not originate from the tip leakage flow, forms in the endwall region within the blade passage and exits the passage near midpitch. Mixing of the leakage vortex with primary flow downstream of the rotor at both design and part speed conditions is also discussed.
Investigation of Transonic Wake Dynamics for Mechanically Deployable Entry Systems
Stern, Eric; Barnhardt, Michael; Venkatapathy, Ethiraj; Candler, Graham; Prabhu, Dinesh
2012-01-01
A numerical investigation of transonic flow around a mechanically deployable entry system being considered for a robotic mission to Venus has been performed, and preliminary results are reported. The flow around a conceptual representation of the vehicle geometry was simulated at discrete points along a ballistic trajectory using Detached Eddy Simulation (DES). The trajectory points selected span the low supersonic to transonic regimes with freestream Mach numbers from 1:5 to 0:8, and freestream Reynolds numbers (based on diameter) between 2:09 x 10(exp 6) and 2:93 x 10(exp 6). Additionally, the Mach 0:8 case was simulated at angles of attack between 0 and 5 . Static aerodynamic coefficients obtained from the data show qualitative agreement with data from 70deg sphere-cone wind tunnel tests performed for the Viking program. Finally, the effect of choices of models and numerical algorithms is addressed by comparing the DES results to those using a Reynolds Averaged Navier-Stokes (RANS) model, as well as to results using a more dissipative numerical scheme.
Directory of Open Access Journals (Sweden)
Yaping Ju
2016-05-01
Full Text Available The Monte Carlo simulation method for turbomachinery uncertainty analysis often requires performing a huge number of simulations, the computational cost of which can be greatly alleviated with the help of metamodeling techniques. An intensive comparative study was performed on the approximation performance of three prospective artificial intelligence metamodels, that is, artificial neural network, radial basis function, and support vector regression. The genetic algorithm was used to optimize the predetermined parameters of each metamodel for the sake of a fair comparison. Through testing on 10 nonlinear functions with different problem scales and sample sizes, the genetic algorithm–support vector regression metamodel was found more accurate and robust than the other two counterparts. Accordingly, the genetic algorithm–support vector regression metamodel was selected and combined with the Monte Carlo simulation method for the uncertainty analysis of a wind turbine airfoil under two types of surface roughness uncertainties. The results show that the genetic algorithm–support vector regression metamodel can capture well the uncertainty propagation from the surface roughness to the airfoil aerodynamic performance. This work is useful to the application of metamodeling techniques in the robust design optimization of turbomachinery.
Luo, Pan; Zhang, Xingwei; Huang, Panpan; Xie, Lingwang
2017-10-01
The aim of this study is to investigate the aerodynamic characteristics of a flapping airfoil in the adjustment stage between two specific flight patterns during the forward flight. Four flapping movement models in adjustment stage are firstly established by using the multi-objective optimization algorithm. Then, a numerical experiment is carried out by using finite volume method to solve the two-dimensional time-dependent incompressible Navier-Stokes equations. The attack angles are selected from -5° to 7.5° with an increase of 2.5°. The results are systematically analyzed and special attention is paid to the corresponding changes of aerodynamic forces, vortex shedding mechanism in the wake structure and thrust efficiency. Present results show that output aerodynamic performance of flapping airfoil can be improved by the increasement of amplitude and frequency in the flapping adjustment stage, which further validates and complements previous studies. Moreover, it is also show that the manner using multi-objective optimization algorithm to generate a movement model in adjustment stage, to connect other two specific plunging motions, is a feasible and effective method. Current study is dedicated to providing some helpful references for the design and control of artificial flapping wing air vehicles.
Influence of transition on steady and unsteady wind-turbine airfoil aerodynamics
Paterson, Eric; Lavely, Adam; Vijayakumar, Ganesh; Brasseur, James
2011-11-01
Laminar-flow airfoils for large stall-regulated horizontal-axis wind turbines are designed to achieve a restrained maximum lift coefficient and a broad laminar low- drag bucket under steady flow conditions and at specific Reynolds numbers. Blind- comparisons of the 2000 NREL Unsteady Aerodynamics Experiment showed large discrepancies and illustrated the need for improved physics modeling. We have studied the S809 airfoil under static and dynamic (ramp-up, ramp-down, and oscillatory) conditions, using the four-equation transition model of Langtry and Menter (2009), which has been implemented as a library accessible by an OpenFOAM RANS solver. Model validation is performed using surface-pressure and lift/drag data from U. Glasgow (2009) and OSU (1995) wind tunnel experiments. Performance of the transition model is assessed by analyzing integrated performance metrics, as well as detailed surface pressure and pressure gradient, wall-shear stress, and boundary-layer profiles and separation points. Demonstration of model performance in the light- and deep-stall regimes of dynamic stall is an important step in reducing uncertainties in full 3D simulations of turbines operating in the atmospheric boundary layer. Supported by NSF Grant 0933647.
Transonic Wind Tunnel Tests on Tonic Mk 5-2: Exercise Yakkata Operational Vehicle
National Research Council Canada - National Science Library
Link, Yoel
2000-01-01
... 0.53 m Transonic Wind Tunnel. This detailed investigation showed that the modifications made in converting Tonic Mk 5-1 to Tonic Mk 5-2 did not have any significant adverse effect on the aerodynamic characteristics...
Analysis of Limit Cycle Oscillation/Transonic High Alpha Flow Visualization. Part 1: Discussion
National Research Council Canada - National Science Library
Cunningham, Atlee M
1998-01-01
...) at low alpha conditions typical of transonic LCO flows with and without tip stores. Laser light sheet/water vapor techniques were used to illuminate the flows, and video recording was used to obtain the data...
National Research Council Canada - National Science Library
Cunningham, Atlee M
1998-01-01
...) at low alpha conditions typical of transonic LCO flows with and without tip stores. Laser light sheet/water vapor techniques were used to illuminate the flows, and video recording was used to obtain the data...
Analysis of Limit Cycle Oscillation/Transonic High Alpha Flow Visualization
National Research Council Canada - National Science Library
Cunningham, Atlee M
1997-01-01
...) at low alpha condition typical of transonic LCO flows with and without tip stores. Laser light sheet/water vapor techniques were used to illuminate the flows, and video recording was used to obtain the data...
National Research Council Canada - National Science Library
Cunningham, Atlee M
1998-01-01
...) at low alpha conditions typical of transonic LCO flows with and without tip stores. Laser light sheet/water vapor techniques were used to illuminate the flows, and video recording was used to obtain the data...
Experimental verification of the new RISOe-A1 airfoil family for wind turbines
Energy Technology Data Exchange (ETDEWEB)
Dahl, K.S.; Fuglsang, P.; Antoniou, I. [Risoe National Lab., Roskilde (Denmark)
1999-03-01
This paper concerns the experimental verification of a new airfoil family for wind turbines. The family consist of airfoils in the relative thickness range from 15% to 30%. Three airfoils, Risoe-A1-18, Risoe-A1-21, and Risoe-A1-24 were tested in a wind tunnel. The verification consisted of both static and dynamic measurements. Here, the static results are presented for a Reynolds number of 1.6x10{sup 6} for the following airfoil configurations: smooth surface (all three airfoils) and Risoe-A1-24 mounted with leading edge roughness, vortex generators, and Gurney-flaps, respectively. All three airfoils have constant lift curve slope and almost constant drag coefficient until the maximum lift coefficient of about 1.4 is reached. The experimental results are compared with corresponding computational from the general purpose flow solver, EllipSys2D, showing good agreement. (au)
Ai, Qing; Kamliya Jawahar, Hasan; Azarpeyvand, Mahdi
2016-01-01
The aerodynamic performance and wake development of a NACA 0012 airfoil fitted with morphing trailing edges were studied using experimental and computational techniques. The NACA 0012 airfoil was tested with morphing trailing edges having various camber profiles with the same trailing edge tip deflection. The aerodynamic force measurements for the airfoil were carried out for a wide range of chord-based Reynolds number and angles of attack with trailing edge deflection angle of β= 5◦ and 10◦....
Airfoil lance apparatus for homogeneous humidification and sorbent dispersion in a gas stream
Myers, Robert B.; Yagiela, Anthony S.
1990-12-25
An apparatus for spraying an atomized mixture into a gas stream comprises a stream line airfoil member having a large radius leading edge and a small radius trailing edge. A nozzle assembly pierces the trailing edge of the airfoil member and is concentrically surrounded by a nacelle which directs shielding gas from the interior of the airfoil member around the nozzle assembly. Flowable medium to be atomized and atomizing gas for atomizing the medium are supplied in concentric conduits to the nozzle. A plurality of nozzles each surrounded by a nacelle are spaced along the trailing edge of the airfoil member.
NUMERICAL INVESTIGATION OF TWO ELEMENT CAMBER MORPHING AIRFOIL IN LOW REYNOLDS NUMBER FLOWS
RAJESH SENTHIL KUMAR T.; V. SIVAKUMAR; BALAJEE RAMAKRISHNANANDA; ARJHUN A.K, SURIYAPANDIYAN
2017-01-01
Aerodynamic performance of a two-element camber morphing airfoil was investigated at low Reynolds number using the transient SST model in ANSYS FLUENT 14.0 and eN method in XFLR5. The two-element camber morphing concept was employed to morph the baseline airfoil into another airfoil by altering the orientation of mean-line at 35% of the chord to achieve better aerodynamic efficiency. NACA 0012 was selected as baseline airfoil. NACA 23012 was chosen as the test case as it has the camber-line s...
A Computational Modeling Mystery Involving Airfoil Trailing Edge Treatments
Choo, Yeunun; Epps, Brenden
2015-11-01
In a curious result, Fairman (2002) observed that steady RANS calculations predicted larger lift than the experimentally-measured data for six different airfoils with non-traditional trailing edge treatments, whereas the time average of unsteady RANS calculations matched the experiments almost exactly. Are these results reproducible? If so, is the difference between steady and unsteady RANS calculations a numerical artifact, or is there a physical explanation? The goals of this project are to solve this thirteen year old mystery and further to model viscous/load coupling for airfoils with non-traditional trailing edges. These include cupped, beveled, and blunt trailing edges, which are common anti-singing treatments for marine propeller sections. In this talk, we present steady and unsteady RANS calculations (ANSYS Fluent) with careful attention paid to the possible effects of asymmetric unsteady vortex shedding and the modeling of turbulence anisotropy. The effects of non-traditional trailing edge treatments are visualized and explained.
Airfoil Trailing Edge Noise Generation and Its Surface Pressure Fluctuation
DEFF Research Database (Denmark)
Zhu, Wei Jun; Shen, Wen Zhong
2015-01-01
where the time history pressure data are recorded by the surface pressure microphones. After the flow-field is stabilized, the generated noise from the airfoil Trailing Edge (TE) is predicted using the acoustic analogy solver, where the results from LES are the input. It is found that there is a strong......In the present work, Large Eddy Simulation (LES) of turbulent flows over a NACA 0015 airfoil is performed. The purpose of such numerical study is to relate the aerodynamic surface pressure with the noise generation. The results from LES are validated against detailed surface pressure measurements...... relation between TE noise and the aerodynamic pressure. The results of power spectrum density show that the fluctuation of aerodynamic pressure is responsible for noise generation....
Drag Reduction of an Airfoil Using Deep Learning
Jiang, Chiyu; Sun, Anzhu; Marcus, Philip
2017-11-01
We reduced the drag of a 2D airfoil by starting with a NACA-0012 airfoil and used deep learning methods. We created a database which consists of simulations of 2D external flow over randomly generated shapes. We then developed a machine learning framework for external flow field inference given input shapes. Past work which utilized machine learning in Computational Fluid Dynamics focused on estimations of specific flow parameters, but this work is novel in the inference of entire flow fields. We further showed that learned flow patterns are transferable to cases that share certain similarities. This study illustrates the prospects of deeper integration of data-based modeling into current CFD simulation frameworks for faster flow inference and more accurate flow modeling.
Numerical simulation of a wind turbine airfoil : part 1
Energy Technology Data Exchange (ETDEWEB)
Ramdenee, D.; Minea, I.S.; Tardiff d' Hamonville, T.; Illinca, A. [Quebec Univ., Rimouski, PQ (Canada). Laboratoire de Recherche en Energie Eolienne
2010-07-01
This 2-part study used computational fluid dynamics (CFD) to identify and model the aerodynamic and aeroelastic phenomena around wind turbine blades. The aim of the study was to better understand the mechanisms surrounding unsteady flow-structure interactions. Aerodynamic and elastic models were coupled using an ANSYS multi-domain program to simulate the aeroelastic divergence of a typical section airfoil with a single rotational structural degree of freedom. Solvers were used to realize a sequence of multi-domain time steps and coupling iterations between time steps. Each element of the airfoil was divided into interpolation faces which were then transformed into 2-D polygons. An intersection process was used to create a large number of control surfaces that were used to study interactions between the structural and fluid domains. The calculations were used to determine the divergence speed and Eigen modes of vibration. A literature review was also included. 19 refs., 7 figs.
Numerical simulations of unsteady transonic flow in diffusers
Liou, M.-S.; Coakley, T. J.
1982-01-01
Forced and naturally occurring, self-sustaining oscillations of transonic flows in two-dimensional diffusers were computed using MacCormack's hybrid method. Depending upon the shock strengths and the area ratios, the flow was fully attached or separated by either the shock or the adverse pressure gradient associated with the enlarging diffuser area. In the case of forced oscillations, a sinusoidal plane pressure wave at frequency 300 Hz was prescribed at the exit. A sufficiently large amount of data were acquired and Fourier analyzed. The distrbutions of time-mean pressures, the power spectral density, and the amplitude with phase angle along the top wall and in the core region were determined. Comparison with experimental results for the forced oscillation generally gave very good agreement; some success was achieved for the case of self-sustaining oscillation despite substantial three-dimensionality in the test. An observation of the sequence of self-sustaining oscillations was given.
Aerodynamics of Airfoils Subject to Three-Dimensional Periodic Gusts.
1983-08-31
and computational procedures to calculate the unsteady forces acting upon airfoils of arbitrary shape subject to three-dimensional gust disturbances...However the mathenatical formulation which has evolved from our analytical work can also be applied under certain conditions to study the changes in...check the validity of our computation scheme two sets of comparisons were carried out. First we considered a two-dimensional gust with transverse and
Airfoil Self-Noise - Investigation with Particle Image Velocimetry
Pröbsting, S.
2015-01-01
Noise generated aerodynamically by the airflow over a lifting surface is often of concern for applications as diverse as air and ground transportation, heating, ventilation, air-conditioning systems, and wind turbines. The thesis describes the application of advanced optical flow measurements techniques for the visualization and description of the sources of sound on airfoils. These measurement techniques include high-speed stereoscopic and tomographic Particle Image Velocimetry (PIV) togethe...
Optimization Approach on Flapping Aerodynamic Characteristics of Corrugated Airfoil
Wei-Hsin Sun; Jr-Ming Miao; Chang-Hsien Tai; Chien-Chun Hung
2011-01-01
The development of biomimetic micro-aerial-vehicles (MAVs) with flapping wings is the future trend in military/domestic field. The successful flight of MAVs is strongly related to the understanding of unsteady aerodynamic performance of low Reynolds number airfoils under dynamic flapping motion. This study explored the effects of flapping frequency, stroke amplitude, and the inclined angle of stroke plane on lift force and thrust force of a bio-inspiration corrugated airf...
Numerical Wake Visualization for Airfoils Undergoing Forced and Aeroelastic Motions
Jones, K.D.; Center, K.B.
1996-01-01
AIAA Paper No. 96-0055, 34th AIAA Aerospace Sciences Meeting, Reno, Nevada, Jan. 1996. A virtual wind tunnel is developed by combining a fast, time-stepping flow solver with an interactive animation interface. Inviscid, incompressible flow solutions are provided by an unsteady, potential-flow code with arbitrary airfoils undergoing forced or aeroelastic motions. Aeroelastic response is predicted by a two-degree-of-freedom spring/mass system modeling the structural dynamics of a flexible wi...
RANS Simulations of Aerodynamic Performance of NACA 0015 Flapped Airfoil
Directory of Open Access Journals (Sweden)
Sohaib Obeid
2017-01-01
Full Text Available An analysis of 2D subsonic flow over an NACA 0015 airfoil with a 30% trailing edge flap at a constant Reynolds number of 106 for various incidence angles and a range of flap deflections is presented. The steady-state governing equations of continuity and momentum conservation are solved combined with the realizable k-ε turbulence model using the ANSYS-Fluent code (Version 13.7, ANSYS, Inc., Canonsburg, PA, USA. The primary objective of the study is to provide a comprehensive understanding of flow characteristics around the NACA 0015 airfoil as a function of the angle of attack and flap deflection at Re = 106 using the realizable k-ε turbulence model. The results are validated through comparison of the predictions with the free field experimental measurements. Consistent with the experimental observations, the numerical results show that increased flap deflections increase the maximum lift coefficient, move the zero-lift angle of attack (AoA to a more negative value, decrease the stall AoA, while the slope of the lift curve remains unchanged and the curve just shifts upwards. In addition, the numerical simulations provide limits for lift increment Δ C l and Cl, max values to be 1.1 and 2.2, respectively, obtained at a flap deflection of 50°. This investigation demonstrates that the realizable k-ε turbulence model is capable of predicting flow features over an airfoil with and without flap deflections with reasonable accuracy.
CFD Study of NACA 0018 Airfoil with Flow Control
Eggert, Christopher A.; Rumsey, Christopher L.
2017-01-01
The abilities of two different Reynolds-Averaged Navier-Stokes codes to predict the effects of an active flow control device are evaluated. The flow control device consists of a blowing slot located on the upper surface of an NACA 0018 airfoil, near the leading edge. A second blowing slot present on the airfoil near mid-chord is not evaluated here. Experimental results from a wind tunnel test show that a slot blowing with high momentum coefficient will increase the lift of the airfoil (compared to no blowing) and delay flow separation. A slot with low momentum coefficient will decrease the lift and induce separation even at low angles of attack. Two codes, CFL3D and FUN3D, are used in two-dimensional computations along with several different turbulence models. Two of these produced reasonable results for this flow, when run fully turbulent. A more advanced transition model failed to predict reasonable results, but warrants further study using different inputs. Including inviscid upper and lower tunnel walls in the simulations was found to be important in obtaining pressure distributions and lift coefficients that best matched experimental data. A limited number of three-dimensional computations were also performed.
Flow control at low Reynolds numbers using periodic airfoil morphing
Jones, Gareth; Santer, Matthew; Papadakis, George; Bouremel, Yann; Debiasi, Marco; Imperial-NUS Joint PhD Collaboration
2014-11-01
The performance of airfoils operating at low Reynolds numbers is known to suffer from flow separation even at low angles of attack as a result of their boundary layers remaining laminar. The lack of mixing---a characteristic of turbulent boundary layers---leaves laminar boundary layers with insufficient energy to overcome the adverse pressure gradient that occurs in the pressure recovery region. This study looks at periodic surface morphing as an active flow control technique for airfoils in such a flight regime. It was discovered that at sufficiently high frequencies an oscillating surface is capable of not only reducing the size of the separated region---and consequently significantly reducing drag whilst simultaneously increasing lift---but it is also capable of delaying stall and as a result increasing CLmax. Furthermore, by bonding Macro Fiber Composite actuators (MFCs) to the underside of an airfoil skin and driving them with a sinusoidal frequency, it is shown that this control technique can be practically implemented in a lightweight, energy efficient way. Imperial-NUS Joint Ph.D. Programme.
Low-Reynolds number compressible flow around a triangular airfoil
Munday, Phillip; Taira, Kunihiko; Suwa, Tetsuya; Numata, Daiju; Asai, Keisuke
2013-11-01
We report on the combined numerical and experimental effort to analyze the nonlinear aerodynamics of a triangular airfoil in low-Reynolds number compressible flow that is representative of wings on future Martian air vehicles. The flow field around this airfoil is examined for a wide range of angles of attack and Mach numbers with three-dimensional direct numerical simulations at Re = 3000 . Companion experiments are conducted in a unique Martian wind tunnel that is placed in a vacuum chamber to simulate the Martian atmosphere. Computational findings are compared with pressure sensitive paint and direct force measurements and are found to be in agreement. The separated flow from the leading edge is found to form a large leading-edge vortex that sits directly above the apex of the airfoil and provides enhanced lift at post stall angles of attack. For higher subsonic flows, the vortical structures elongate in the streamwise direction resulting in reduced lift enhancement. We also observe that the onset of spanwise instability for higher angles of attack is delayed at lower Mach numbers. Currently at Mitsubishi Heavy Industries, Ltd., Nagasaki.
Simulasi Numerik Dynamic Stall Pada Airfoil Yang Berosilasi
Directory of Open Access Journals (Sweden)
Galih S.T.A. Bangga
2012-09-01
Full Text Available Kebutuhan analisa pada sudu helikopter, kompresor, kincir angin dan struktur streamline lainya yang beroperasi pada angle of attack yang tinggi dan melibatkan instationary effects yang disebut dynamic stall menjadi semakin penting. Fenomena ini ditandai dengan naiknya dynamic lift melewati static lift maksimum pada critical static stall angle, vortex yang terbentuk pada leading edge mengakibatkan naiknya suction contribution yang kemudian terkonveksi sepanjang permukaan hingga mencapai trailling edge diikuti terbentuknya trailling edge vortex yang menunjukkan terjadinya lift stall. Fenomena ini sangat berbahaya terhadap struktur airfoil itu sendiri. Secara umum, beban fatique yang ditimbulkan oleh adanya efek histerisis karena fluktuasi gaya lift akibat induksi vibrasi lebih besar dibandingkan kondisi statis. Simulasi numerik dilakukan secara 2D dengan menggunakan profil Boeing-Vertol V23010-1.58 pada α0 = 14.92°. Standard-kω dan SST-kω digunakan sebagai URANS turbulence modelling. Model osilasi dari airfoil disusun dalam suatu user defined function (UDF. Gerakan meshing beserta airfoil diakomodasi dengan menggunakan dynamic mesh approach. Simulasi numerik menunjukkan bahwa, model SST-kω menunjukkan performa yang lebih baik dibandingkan dengan Standard-kω. Fenomena travelling vortex yang terjadi mampu ditangkap dengan baik, meski pada angle of attack yang tinggi URANS turbulence model gagal memprediksikan fenomena yang terjadi karena dominasi efek 3D.
Kolb, Mark A.
1990-01-01
Viewgraphs on Rubber Airplane: Constraint-based Component-Modeling for Knowledge Representation in Computer Aided Conceptual Design are presented. Topics covered include: computer aided design; object oriented programming; airfoil design; surveillance aircraft; commercial aircraft; aircraft design; and launch vehicles.
Development of a Fowler flap system for a high performance general aviation airfoil
Wentz, W. H., Jr.; Seetharam, H. C.
1974-01-01
A two-dimensional wind-tunnel evaluation of two Fowler flap configurations on the new GA(W)-1 airfoil was conducted. One configuration used a computer-designed 29-percent chord Fowler flap. The second configuration was modified to have increased Fowler action with a 30-percent chord flap. Force, pressure, and flow-visualization data were obtained at Reynolds numbers of 2.2 million to 2.9 million. Optimum slot geometry and performance were found to be close to computer predictions. A C sub L max of 3.8 was achieved. Optimum flap deflection, slot gap, and flap overlap are presented as functions of C sub L. Tests were made with the lower surface cusp filled in to show the performance penalties that result. Some data on the effects of adding vortex generators and hinged-plate spoilers were obtained.
Test results of NREL 10M, special-purpose family of thin airfoils
Energy Technology Data Exchange (ETDEWEB)
Starcher, K.L.; Nelson, V.C.; Wei, Jun [West Texas A& M Univ., Canyon, TX (United States)
1996-12-31
Two 25 kW Carter Wind Systems were tested to determine performance differences between production blades and rotors with NREL Special Purpose Thin Airfoils. Blade design, mold preparation, blade production, and testing were conducted. Design tools were created for computer modeling of the blade. The blades had the same twist, taper, and length as production blades. Flap natural frequency was adjusted to be as similar as possible between rotors, as was blade mass, blade center of gravity and rotor moment of inertia. Data collected were; wind speed at hub height, blade root flap & edgewise loads, main shaft torque, azimuth position, teeter angle, yaw angle and electrical power. These data were collected at 128 Hertz for data sets of eight seconds. This data set was then written to hard disk and the cycle repeated resulting in a file containing five and one half minutes of data. A data run consisted of; preflight checkout/warm-up of equipment, preflight calibration/verification of all sensors on both turbines, collection of five files of data (about thirty minutes of data), post flight calibration/verification of sensors. During this high speed data collection period there were a total of twenty-four data runs collected. Data were collected for wind speeds in the range about 7, 10 and 13 m/s. A data matrix was filled for clean, medium and heavy surface roughness. Baseline power curves, parametric pitch variation runs to establish testing pitch settings, high speed data collection runs with and without applied surface roughness were completed and analyzed. Data were compared using simple arithmetic mean, Fast Fourier Transform (FFT) analysis, rainflow counting algorithms and wavelet analysis. The NREL airfoils showed much less sensitivity to surface roughness. There were minimal root bending load differences. Annual energy production during long term operation is being determined. 9 refs., 7 figs.
A Numerical Study of Aerodynamic Performance and Noise of a Bionic Airfoil Based on Owl Wing
Directory of Open Access Journals (Sweden)
Xiaomin Liu
2014-08-01
Full Text Available Noise reduction and efficiency enhancement are the two important directions in the development of the multiblade centrifugal fan. In this study, we attempt to develop a bionic airfoil based on the owl wing and investigate its aerodynamic performance and noise-reduction mechanism at the relatively low Reynolds number. Firstly, according to the geometric characteristics of the owl wing, a bionic airfoil is constructed as the object of study at Reynolds number of 12,300. Secondly, the large eddy simulation (LES with the Smagorinsky model is adopted to numerically simulate the unsteady flow fields around the bionic airfoil and the standard NACA0006 airfoil. And then, the acoustic sources are extracted from the unsteady flow field data, and the Ffowcs Williams-Hawkings (FW-H equation based on Lighthill's acoustic theory is solved to predict the propagation of these acoustic sources. The numerical results show that the lift-to-drag ratio of bionic airfoil is higher than that of the traditional NACA 0006 airfoil because of its deeply concave lower surface geometry. Finally, the sound field of the bionic airfoil is analyzed in detail. The distribution of the A-weighted sound pressure levels, the scaled directivity of the sound, and the distribution of dP/dt on the airfoil surface are provided so that the characteristics of the acoustic sources could be revealed.
Silva, Walter A.
1993-01-01
A methodology for modeling nonlinear unsteady aerodynamic responses, for subsequent use in aeroservoelastic analysis and design, using the Volterra-Wiener theory of nonlinear systems is presented. The methodology is extended to predict nonlinear unsteady aerodynamic responses of arbitrary frequency. The Volterra-Wiener theory uses multidimensional convolution integrals to predict the response of nonlinear systems to arbitrary inputs. The CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) code is used to generate linear and nonlinear unit impulse responses that correspond to each of the integrals for a rectangular wing with a NACA 0012 section with pitch and plunge degrees of freedom. The computed kernels then are used to predict linear and nonlinear unsteady aerodynamic responses via convolution and compared to responses obtained using the CAP-TSD code directly. The results indicate that the approach can be used to predict linear unsteady aerodynamic responses exactly for any input amplitude or frequency at a significant cost savings. Convolution of the nonlinear terms results in nonlinear unsteady aerodynamic responses that compare reasonably well with those computed using the CAP-TSD code directly but at significant computational cost savings.
Turbine Design to Mitigate Forcing (POSTPRINT)
2012-09-01
blowing from the pressure side of the downstream stationary airfoil row. In addition, the effects of downstream vane asymmetric spacing and vane-to- vane...steady blowing from the pressure side of the downstream stationary airfoil row. In addition, the effects of downstream vane asymmetric spacing and vane...give the final geometry of the experimental turbine. This resulted in what is often called a “balanced shock” blade design. 30 Approved for
Smeltzer, D. B.; Sorensen, N. E.
1972-01-01
A 38.8-cm (15.28-in.) capture diameter model of a mixed-compression axisymmetric inlet system with a translating cowl was designed and tested. The internal contours, designed for Mach number 2.65, provided a throat area of 59 percent of the capture area when the cowl was retracted for transonic operation. Other model features included a boundary-layer removal system, vortex generators, an engine airflow bypass system, cowl support struts, and rotating rakes at the engine face. All tunnel testing was conducted at a tunnel total pressure of about 1 atm (a unit Reynolds number of about 8.53 million/m at Mach number 2.65) at angles of attack from 0 deg to 4 deg. Results for the following were obtained: total-pressure recovery and distortion at the engine face as a function of bleed mass-flow ratio, the effect of bleed and vortex generator configurations on pressure recovery and distortion, inlet tolerance to unstart due to changes in angle of attack or Mach number, surface pressure distributions, boundary-layer profiles, and transonic additive drag. At Mach number 2.65 and with the best bleed configurations, maximum total pressure recovery at the engine face ranged from 91 to 94.5 percent with bleed mass-flow ratios from 4 to 9 percent, respectively, and total-pressure distortion was less than 10 percent. At off-design supersonic Mach numbers above 1.70, maximum total-pressure recoveries and corresponding bleed mass flows were about the same as at Mach number 2.65, with about 10 to 15 percent distortion. In the transonic Mach number range, total pressure recovery was high (above 96 percent) and distortion was low (less than 15 percent) only when the inlet mass-flow ration was reduced 0.02 to 0.06 from the maximum theoretical value (0.590 at Mach number 1.0).
Influences of surface temperature on a low camber airfoil aerodynamic performances
Directory of Open Access Journals (Sweden)
Valeriu DRAGAN
2016-03-01
Full Text Available The current note refers to the comparison between a NACA 2510 airfoil with adiabatic walls and the same airfoil with heated patches. Both suction and pressure sides were divided into two regions covering the leading edge (L.E. and trailing edge (T.E.. A RANS method sensitivity test has been performed in the preliminary stage while for the extended 3D cases a DES-SST approach was used. Results indicate that surface temperature distribution influences the aerodynamics of the airfoil, in particular the viscous drag component but also the lift of the airfoil. Moreover, the influence depends not only on the surface temperature but also on the positioning of the heated surfaces, particularly in the case of pressure lift and drag. Further work will be needed to optimize the temperature distribution for airfoil with higher camber.
Optimization of bump and blowing to control the flow through a transonic compressor blade cascade
Mazaheri, K.; Khatibirad, S.
2018-03-01
Shock control bump (SCB) and blowing are two flow control methods, used here to improve the aerodynamic performance of transonic compressors. Both methods are applied to a NASA rotor 67 blade section and are optimized to minimize the total pressure loss. A continuous adjoint algorithm is used for multi-point optimization of a SCB to improve the aerodynamic performance of the rotor blade section, for a range of operational conditions around its design point. A multi-point and two single-point optimizations are performed in the design and off-design conditions. It is shown that the single-point optimized shapes have the best performance for their respective operating conditions, but the multi-point one has an overall better performance over the whole operating range. An analysis is given regarding how similarly both single- and multi-point optimized SCBs change the wave structure between blade sections resulting in a more favorable flow pattern. Interactions of the SCB with the boundary layer and the wave structure, and its effects on the separation regions are also studied. We have also introduced the concept of blowing for control of shock wave and boundary-layer interaction. A geometrical model is introduced, and the geometrical and physical parameters of blowing are optimized at the design point. The performance improvements of blowing are compared with the SCB. The physical interactions of SCB with the boundary layer and the shock wave are analyzed. The effects of SCB on the wave structure in the flow domain outside the boundary-layer region are investigated. It is shown that the effects of the blowing mechanism are very similar to the SCB.
Unsteady airfoil flows with application to aeroelastic stability
Energy Technology Data Exchange (ETDEWEB)
Johansen, Jeppe
1999-09-01
The present report describes numerical investigation of two-dimensional unsteady airfoil flows with application to aeroelastic stability. The report is divided in two parts. Part A describes the purely aerodynamic part, while Part B includes the aeroelastic part. In Part A a transition prediction algorithm based on a simplified version of the e{sup n} method is proposed. Laminar Boundary Layer instability data are stored in a database from which stability characteristics can be extracted by interpolation. Input to the database are laminar integral boundary layer parameters. These are computed from an integral boundary layer formulation coupled to a Navier-Stokes flow solver. Five different airfoils are considered at fixed angle of attack, and the flow is computed assuming both fully turbulent and transitional flow and compared with experimental data. Results indicate that using a transition model the drag prediction is improved considerably. Also the lift is slightly improved. At high angles of attack transition will affect leading edge separation which again will affect the overall vortex shedding. If the transition point is not properly predicted this will affect the whole hysteresis curve. The transition model developed in the present work showed more stable predictions compared to the empirical transition model. In Part B a simple three degrees-of-freedom (DOF) structural dynamics model is developed and coupled to the aerodynamics models from Part A. A 2nd order accurate time integration scheme is used to solve the equations of motion. Two airfoils are investigated. The aeroelastic models predict stable conditions well at low angle of attack. But at high angles of attack, and where unstable behaviour is expected, only the Navier-Stokes solver predict correct aeroelastic response. The semi-empirical dynamic stall model does not predict vortex shedding and moment correctly leading to an erroneous aerodynamic damping. (au) 5 tabs.; 55 ills., 52 refs.
Unsteady Thick Airfoil Aerodynamics: Experiments, Computation, and Theory
Strangfeld, C.; Rumsey, C. L.; Mueller-Vahl, H.; Greenblatt, D.; Nayeri, C. N.; Paschereit, C. O.
2015-01-01
An experimental, computational and theoretical investigation was carried out to study the aerodynamic loads acting on a relatively thick NACA 0018 airfoil when subjected to pitching and surging, individually and synchronously. Both pre-stall and post-stall angles of attack were considered. Experiments were carried out in a dedicated unsteady wind tunnel, with large surge amplitudes, and airfoil loads were estimated by means of unsteady surface mounted pressure measurements. Theoretical predictions were based on Theodorsen's and Isaacs' results as well as on the relatively recent generalizations of van der Wall. Both two- and three-dimensional computations were performed on structured grids employing unsteady Reynolds-averaged Navier-Stokes (URANS). For pure surging at pre-stall angles of attack, the correspondence between experiments and theory was satisfactory; this served as a validation of Isaacs theory. Discrepancies were traced to dynamic trailing-edge separation, even at low angles of attack. Excellent correspondence was found between experiments and theory for airfoil pitching as well as combined pitching and surging; the latter appears to be the first clear validation of van der Wall's theoretical results. Although qualitatively similar to experiment at low angles of attack, two-dimensional URANS computations yielded notable errors in the unsteady load effects of pitching, surging and their synchronous combination. The main reason is believed to be that the URANS equations do not resolve wake vorticity (explicitly modeled in the theory) or the resulting rolled-up un- steady flow structures because high values of eddy viscosity tend to \\smear" the wake. At post-stall angles, three-dimensional computations illustrated the importance of modeling the tunnel side walls.
A dynamic stall model for airfoils with deformable trailing edges
DEFF Research Database (Denmark)
Andersen, Peter Bjørn; Gaunaa, Mac; Bak, Christian
2009-01-01
The present work contains an extension of the Beddoes-Leishman-type dynamic stall model. In this work, a deformable trailing-edge flap has been added to the dynamic stall model. The model predicts the unsteady aerodynamic forces and moments on an airfoil section undergoing arbitrary motion in hea...... for the attached flow region and Hansen et al. The model is compared qualitatively to wind tunnel measurements of a Riso/ B1-18 blade section equipped with deformable trailing-edge flap devices in the form of piezoelectric devices. Copyright © 2009 John Wiley & Sons, Ltd....
Development of heat flux sensors for turbine airfoils
Atkinson, William H.; Cyr, Marcia A.; Strange, Richard R.
1985-10-01
The objectives of this program are to develop heat flux sensors suitable for installation in hot section airfoils of advanced aircraft turbine engines and to experimentally verify the operation of these heat flux sensors in a cylinder in a cross flow experiment. Embedded thermocouple and Gardon gauge sensors were developed and fabricated into both blades and vanes. These were then calibrated using a quartz lamp bank heat source and finally subjected to thermal cycle and thermal soak testing. These sensors were also fabricated into cylindrical test pieces and tested in a burner exhaust to verify heat flux measurements produced by these sensors. The results of the cylinder in cross flow tests are given.
Iced airfoil separation bubble measurements by particle image velocimetry
Jacobs, Jason J.
Not long after the birth of aviation, pilots began to recognize the dangers posed by aircraft icing. Since that time, research has improved the awareness of this problem and the scientific understanding of the associated aerodynamic impacts, however, few studies have involved detailed, quantitative, flowfield measurements. For this reason, the current investigation was conducted in which high spatial-resolution flowfield measurements were acquired of a NACA 0012 airfoil with two- and three-dimensional, simulated, leading-edge, horn-ice accretions utilizing particle image velocimetry (PIV). These measurements complemented existing iced airfoil performance measurements, revealed previously unknown details regarding the structure and behavior of these flowfields, and could potentially facilitate the development and improvement of computational schemes used to predict largely separated flows, including that of an iced airfoil near stall. Previous iced airfoil investigations have demonstrated somewhat reduced aerodynamic penalties resulting from a three-dimensional ice simulation, compared to those of a two-dimensional ice simulation of a representative cross section. Correspondingly, the current measurements revealed accelerated transition of the separated shear layer emanating from a three-dimensional ice simulation and therefore enhanced pressure recovery and reduced mean separation bubble length, each relative to the flowfield of a representative two-dimensional ice simulation. These effects appeared to result from the quasi-steady distribution of discrete, streamwise vortices which aided the turbulent entrainment of fluid from the recirculation region of the three-dimensional ice simulation separation bubble flowfield. These vortices were generated by a streamwise-vortex instability excited by roughness along the three-dimensional ice simulation and produced spanwise-cell structures throughout this flowfield, as well as significant spanwise variation in peak
HiMAT aerodynamic design and flight test experience
Matheny, N. W.; Panageas, G. N.
1981-01-01
Consideration is given to the design phase of the highly maneuverable aircraft technology program. Design objectives are examined, noting full-scale design and the remotely piloted research vehicle. Attention is given to subsonic, transonic, and supersonic design. Design results are discussed with reference to aerodynamic efficiency, aeroelastic tailoring, and the flight test program.
Three-dimensional flows in a transonic compressor rotor
Reid, Lonnie; Celestina, Mark L.; Dewitt, Kenneth; Keith, Theo
1991-01-01
This study involves an experimental and numerical investigation of the three-dimensional flows in a transonic compressor rotor. A variety of data which could be used, in a complementary fashion, to validate/calibrate the computational fluid dynamics turbomachinery code and improve understanding of the flow physics, were acquired. Detailed radial survey data which consisted of total pressure, total temperature, static pressure and flow angle were obtained at stations upstream and downstream of the rotor blade. Detailed velocity and turbulence profiles were obtained upstream of the rotor and used as the upstream boundary conditions for the numerical analysis. Calibrated flush-mounted hot film probes were used to measure wall shear stress on the hub and casing walls upstream of the rotor. The blade-to-blade shear-stress angle distributions were obtained at two axial locations on the rotor casing, using flush-mounted hot film probes. A numerical analysis conducted using a three-dimensional Navier-Stokes code was compared with the experimental results.
Contribution to the study of unsteady condensation in transonic flow
International Nuclear Information System (INIS)
Collignan, B.; Laali, A.R.
1993-12-01
The aim of this thesis is the study of transonic steam flows with condensation, especially at high pressure. This study includes a numerical part an experimental one. The modelling has consisted of introducing a spontaneous condensation model in a one-dimensional Euler code using steam-water thermodynamic tables. Calculations, performed with this code, are in good agreement with experimental results at low pressure. The experimental study has been undertaken on a high pressure experimental loop installed at the Bugey nuclear power plant. We have studied steam flows in nozzles. The results obtained show that a partial heterogeneous condensation occurs in these flows. This proportion is stronger if the expansion rate of the flow is low and if the inlet pressure is high. However, a correction factor is obtained for high pressure nucleation rate model from experimental results. No unsteady condensation has been observed for flows between 15 bars and 50 bars with the steam available at Bugey power plant. (authors). figs., 71 refs., 6 annexes
Taherian, Gholamhossein; Nili-Ahmadabadi, Mahdi; Karimi, Mohammad Hassan; Tavakoli, Mohammad Reza
2017-01-01
In this study, the effect of cutting the end of a thick airfoil and adding a cavity on its flow pattern is studied experimentally using PIV technique. First, by cutting 30% chord length of the Riso airfoil, a thick blunt trialing-edge airfoil is generated. The velocity field around the original airfoil and the new airfoil is measured by PIV technique and compared with each other. Then, adding two parallel plates to the end of the new airfoil forms the desired cavity. Continuous measurement of unsteady flow velocity over the Riso airfoil with thick blunt trailing edge and base cavity is the most important innovation of this research. The results show that cutting off the end of the airfoil decreases the wake region behind the airfoil, when separation occurs. Moreover, adding a cavity to the end of the thickened airfoil causes an increase in momentum and a further decrease in the wake behind the trailing edge that leads to a drag reduction in comparison with the thickened airfoil without cavity. Furthermore, using cavity decreases the Strouhal number and vortex shedding frequency.
CFD Predictions for Transonic Performance of the ERA Hybrid Wing-Body Configuration
Deere, Karen A.; Luckring, James M.; McMillin, S. Naomi; Flamm, Jeffrey D.; Roman, Dino
2016-01-01
A computational study was performed for a Hybrid Wing Body configuration that was focused at transonic cruise performance conditions. In the absence of experimental data, two fully independent computational fluid dynamics analyses were conducted to add confidence to the estimated transonic performance predictions. The primary analysis was performed by Boeing with the structured overset-mesh code OVERFLOW. The secondary analysis was performed by NASA Langley Research Center with the unstructured-mesh code USM3D. Both analyses were performed at full-scale flight conditions and included three configurations customary to drag buildup and interference analysis: a powered complete configuration, the configuration with the nacelle/pylon removed, and the powered nacelle in isolation. The results in this paper are focused primarily on transonic performance up to cruise and through drag rise. Comparisons between the CFD results were very good despite some minor geometric differences in the two analyses.
Polytropic transonic galactic outflows in a dark matter halo with a central black hole
Igarashi, Asuka; Mori, Masao; Nitta, Shin-ya
2017-09-01
Polytropic transonic solutions of spherically symmetric and steady galactic winds in the gravitational potential of a dark matter halo (DMH) with a supermassive black hole (SMBH) are studied. The solutions are classified in terms of their topological features, and the gravitational potential of the SMBH adds a new branch to the transonic solutions generated by the gravity of the DMH. The topological types of the transonic solutions depend on the mass distribution, the amount of supplied energy, the polytropic index γ and the slope α of the DMH mass distribution. When α becomes larger than a critical value αc, the transonic solution types change dramatically. Further, our model predicts that it is possible for a slowly accelerating outflow to exist, even in quiescent galaxies with small γ. This slowly accelerating outflow differs from those considered in many of the previous studies focusing on supersonic outflows in active star-forming galaxies. In addition, our model indicates that outflows in active star-forming galaxies have only one transonic point in the inner region (˜0.01 kpc). The locus of this transonic point does not strongly depend on γ. We apply the polytropic model incorporating mass flux supplied by stellar components to the Sombrero galaxy, and conclude that it can reproduce the observed gas density and the temperature distribution well. This result differs significantly from the isothermal model, which requires an unrealistically large mass flux. Thus, we conclude that the polytropic model is more realistic than the isothermal model, and that the Sombrero galaxy can have a slowly accelerating outflow.
Airfoil-Shaped Fluid Flow Tool for Use in Making Differential Measurements
England, John Dwight (Inventor); Kelley, Anthony R. (Inventor); Cronise, Raymond J. (Inventor)
2014-01-01
A fluid flow tool includes an airfoil structure and a support arm. The airfoil structure's high-pressure side and low-pressure side are positioned in a conduit by the support arm coupled to the conduit. The high-pressure and low-pressure sides substantially face opposing walls of the conduit. At least one measurement port is formed in the airfoil structure at each of its high-pressure side and low-pressure side. A first manifold, formed in the airfoil structure and in fluid communication with each measurement port so-formed at the high-pressure side, extends through the airfoil structure and support arm to terminate and be accessible at the exterior wall of the conduit. A second manifold, formed in the airfoil structure and in fluid communication with each measurement port so-formed at the low-pressure side, extends through the airfoil structure and support arm to terminate and be accessible at the exterior wall of the conduit.
Parametric analyses for synthetic jet control on separation and stall over rotor airfoil
Directory of Open Access Journals (Sweden)
Zhao Guoqing
2014-10-01
Full Text Available Numerical simulations are performed to investigate the effects of synthetic jet control on separation and stall over rotor airfoils. The preconditioned and unsteady Reynolds-averaged Navier–Stokes equations coupled with a k − ω shear stream transport turbulence model are employed to accomplish the flowfield simulation of rotor airfoils under jet control. Additionally, a velocity boundary condition modeled by a sinusoidal function is developed to fulfill the perturbation effect of periodic jets. The validity of the present CFD procedure is evaluated by the simulated results of an isolated synthetic jet and the jet control case for airfoil NACA0015. Then, parametric analyses are conducted specifically for an OA213 rotor airfoil to investigate the effects of jet parameters (forcing frequency, jet location and momentum coefficient, jet direction, and distribution of jet arrays on the control effect of the aerodynamic characteristics of a rotor airfoil. Preliminary results indicate that the efficiency of jet control can be improved with specific frequencies (the best lift-drag ratio at F+ = 2.0 and jet angles (40° or 75° when the jets are located near the separation point of the rotor airfoil. Furthermore, as a result of a suitable combination of jet arrays, the lift coefficient of the airfoil can be improved by nearly 100%, and the corresponding drag coefficient decreased by 26.5% in comparison with the single point control case.
Experimental investigation of trailing edge noise from stationary and rotating airfoils.
Zajamsek, Branko; Doolan, Con J; Moreau, Danielle J; Fischer, Jeoffrey; Prime, Zebb
2017-05-01
Trailing edge noise from stationary and rotating NACA 0012 airfoils is characterised and compared with a noise prediction based on the semi-empirical Brooks, Pope, and Marcolini (BPM) model. The NACA 0012 is symmetrical airfoil with no camber and 12% thickness to chord length ratio. Acoustic measurements were conducted in an anechoic wind tunnel using a stationary NACA 0012 airfoil at 0° pitch angle. Airfoil self-noise emissions from rotating NACA 0012 airfoils mounted at 0° and 10° pitch angles on a rotor-rig are studied in an anechoic room. The measurements were carried out using microphone arrays for noise localisation and magnitude estimation using beamforming post-processing. Results show good agreement between peak radiating trailing edge noise emissions of stationary and rotating NACA 0012 airfoils in terms of the Strouhal number. Furthermore, it is shown that noise predictions based on the BPM model considering only two dimensional flow effects, are in good agreement with measurements for rotating airfoils, at these particular conditions.
Unsteady 2D potential-flow forces and a thin variable geometry airfoil undergoing arbitrary motion
Energy Technology Data Exchange (ETDEWEB)
Gaunaa, M.
2006-07-15
In this report analytical expressions for the unsteady 2D force distribution on a variable geometry airfoil undergoing arbitrary motion are derived under the assumption of incompressible, irrotational, inviscid flow. The airfoil is represented by its camberline as in classic thin-airfoil theory, and the deflection of the airfoil is given by superposition of chordwise deflection mode shapes. It is shown from the expressions for the forces, that the influence from the shed vorticity in the wake is described by the same time-lag for all chordwise positions on the airfoil. This time-lag term can be approximated using an indicial function approach, making the practical calculation of the aerodynamic response numerically very efficient by use of Duhamel superposition. Furthermore, the indicial function expressions for the time-lag terms are formulated in their equivalent state-space form, allowing for use of the present theory in problems employing the eigenvalue approach, such as stability analysis. The analytical expressions for the forces simplify to all previously known steady and unsteady thin-airfoil solutions. Apart from the obvious applications within active load control/reduction, the current theory can be used for various applications which up to now have been possible only using much more computational costly methods. The propulsive performance of a soft heaving propulsor, and the influence of airfoil camberline elasticity on the flutter limit are two computational examples given in the report that highlight this feature. (au)
On the effect of leading edge blowing on circulation control airfoil aerodynamics
Mclachlan, B. G.
1987-01-01
In the present context the term circulation control is used to denote a method of lift generation that utilizes tangential jet blowing over the upper surface of a rounded trailing edge airfoil to determine the location of the boundary layer separation points, thus setting an effective Kutta condition. At present little information exists on the flow structure generated by circulation control airfoils under leading edge blowing. Consequently, no theoretical methods exist to predict airfoil performance under such conditions. An experimental study of the flow field generated by a two dimensional circulation control airfoil under steady leading and trailing edge blowing was undertaken. The objective was to fundamentally understand the overall flow structure generated and its relation to airfoil performance. Flow visualization was performed to define the overall flow field structure. Measurements of the airfoil forces were also made to provide a correlation of the observed flow field structure to airfoil performance. Preliminary results are presented, specifically on the effect on the flow field structure of leading edge blowing, alone and in conjunction with trailing edge blowing.
Closed-Loop Aerodynamic Flow Control of a Maneuvering Airfoil
Brzozowski, Daniel P.; Culp, John R.; Glezer, Ari
2011-11-01
The unsteady interaction between trailing edge aerodynamic flow control and airfoil motion in pitch and plunge is investigated in wind tunnel experiments using a 2-DOF traverse which enables application of time-dependent external torque and forces by servo motors. The global aerodynamic forces and moments are regulated by controlling vorticity generation and accumulation near the surface using hybrid synthetic jet actuators. The dynamic coupling between the actuation and the time-dependent flow field is characterized using simultaneous force and velocity measurements that are taken phase-locked to the commanded actuation waveform. The effect of the unsteady motion on the model-embedded flow control is assessed in unsteady several maneuvers. Circulation time history that is estimated from a PIV wake survey shows that the entire flow over the airfoil readjusts within about 1.5 TCONV, which is about two orders of magnitude shorter than the characteristic time associated with the controlled maneuver of the wind tunnel model. This illustrates that flow-control actuation can be typically effected on time scales that are commensurate with the flow's convective time scale, and that the maneuver response is primarily limited by the inertia of the platform.
Dynamic Stall in Pitching Airfoils: Aerodynamic Damping and Compressibility Effects
Corke, Thomas C.; Thomas, Flint O.
2015-01-01
Dynamic stall is an incredibly rich fluid dynamics problem that manifests itself on an airfoil during rapid, transient motion in which the angle of incidence surpasses the static stall limit. It is an important element of many manmade and natural flyers, including helicopters and supermaneuverable aircraft, and low-Reynolds number flapping-wing birds and insects. The fluid dynamic attributes that accompany dynamic stall include an eruption of vorticity that organizes into a well-defined dynamic stall vortex and massive excursions in aerodynamic loads that can couple with the airfoil structural dynamics. The dynamic stall process is highly sensitive to surface roughness that can influence turbulent transition and to local compressibility effects that occur at free-stream Mach numbers that are otherwise incompressible. Under some conditions, dynamic stall can result in negative aerodynamic damping that leads to limit-cycle growth of structural vibrations and rapid mechanical failure. The mechanisms leading to negative damping have been a principal interest of recent experiments and analysis. Computational fluid dynamic simulations and low-order models have not been good predictors so far. Large-eddy simulation could be a viable approach although it remains computationally intensive. The topic is technologically important owing to the desire to develop next-generation rotorcraft that employ adaptive rotor dynamic stall control.
Improvement of airfoil trailing edge bluntness noise model
Directory of Open Access Journals (Sweden)
Wei Jun Zhu
2016-02-01
Full Text Available In this article, airfoil trailing edge bluntness noise is investigated using both computational aero-acoustic and semi-empirical approach. For engineering purposes, one of the most commonly used prediction tools for trailing edge noise are based on semi-empirical approaches, for example, the Brooks, Pope, and Marcolini airfoil noise prediction model developed by Brooks, Pope, and Marcolini (NASA Reference Publication 1218, 1989. It was found in previous study that the Brooks, Pope, and Marcolini model tends to over-predict noise at high frequencies. Furthermore, it was observed that this was caused by a lack in the model to predict accurately noise from blunt trailing edges. For more physical understanding of bluntness noise generation, in this study, we also use an advanced in-house developed high-order computational aero-acoustic technique to investigate the details associated with trailing edge bluntness noise. The results from the numerical model form the basis for an improved Brooks, Pope, and Marcolini trailing edge bluntness noise model.
Modeling unsteady forces and pressures on a rapidly pitching airfoil
Schiavone, Nicole K.; Dawson, Scott T. M.; Rowley, Clarence W.; Williams, David R.
2014-11-01
This work develops models to quantify and understand the unsteady aerodynamic forces arising from rapid pitching motion of a NACA0012 airfoil at a Reynolds number of 50 000. The system identification procedure applies a generalized DMD-type algorithm to time-resolved wind tunnel measurements of the lift and drag forces, as well as the pressure at six locations on the suction surface of the airfoil. Models are identified for 5-degree pitch-up and pitch-down maneuvers within the overall range of 0-20 degrees. The identified models can accurately capture the effects of flow separation and leading-edge vortex formation and convection. We demonstrate that switching between different linear models can give accurate prediction of the nonlinear behavior that is present in high-amplitude maneuvers. The models are accurate for a wide-range of motions, including pitch-and-hold, sinusoidal, and pseudo-random pitching maneuvers. Providing the models access to a subset of the measured data channels can allow for improved estimates of the remaining states via the use of a Kalman filter, suggesting that the modeling framework could be useful for aerodynamic control applications. This work was supported by the Air Force Office of Scientific Research, under Award No. FA9550-12-1-0075.
Parametric analyses on dynamic stall control of rotor airfoil via synthetic jet
Directory of Open Access Journals (Sweden)
Qijun ZHAO
2017-12-01
Full Text Available The effects of synthetic jet control on unsteady dynamic stall over rotor airfoil are investigated numerically. A moving-embedded grid method and an Unsteady Reynolds Averaged Navier-Stokes (URANS solver coupled with k-Ï Shear Stress Transport (SST turbulence model are established for predicting the complex flowfields of oscillatory airfoil under jet control. Additionally, a velocity boundary condition modeled by sinusoidal function has been developed to fulfill the perturbation effect of periodic jet. The validity of present CFD method is evaluated by comparisons of the calculated results of baseline dynamic stall case for rotor airfoil and jet control case for VR-7B airfoil with experimental data. Then, parametric analyses are conducted emphatically for an OA212 rotor airfoil to investigate the effects of jet control parameters (jet location, dimensionless frequency, momentum coefficient, jet angle, jet type and dual-jet on dynamic stall characteristics of rotor airfoil. It is demonstrated by the calculated results that efficiency of jet control could be improved with specific momentum coefficient and jet angle when the jet is located near separation point of rotor airfoil. Furthermore, the dual-jet could improve control efficiency more obviously on dynamic stall of rotor airfoil with respect to the unique jet, and the influence laws of dual-jetâs angles and momentum coefficients on control effects are similar to those of the unique jet. Finally, unsteady aerodynamic characteristics of rotor via synthetic jet which is located on the upper surface of rotor blade in forward flight are calculated, and asÂ a result, the aerodynamic characteristics of rotor are improved compared with the baseline. The results indicate that synthetic jet has the capability in improving aerodynamic characteristics of rotor. Keywords: Airfoil, Dynamic stall characteristics, Flow control, Moving-embedded grid methodology, Navier-Stokes equations, Parametric
New airfoil sections for general aviation aircraft. [cruising and flap development tests
Wentz, W. H., Jr.
1973-01-01
A program has been undertaken to develop new airfoil sections suitable for general aviation aircraft, utilizing theoretical and experimental advanced technology developed in recent years primarily for subsonic jet transport and military aircraft. The airfoil development program is one component of the Advanced Technology Light Twin program sponsored by NASA Langley Research Center. Two-dimensional tests of a new airfoil have demonstrated high cruising performance over a fairly wide C sub 1 range, and a C sub 1 max value of 3.69 with Fowler flap and no leading-edge devices. Experimental and theoretical development of additional configurations is under way.
Experimental Study of Tip Vortex Flow from a Periodically Pitched Airfoil Section
Zaman, Khairul; Fagan, Amy; Mankbadi, Mina
2016-01-01
An experimental investigation of tip vortex flow from a NACA0012 airfoil, pitched periodically at various frequencies, is conducted in a low-speed wind tunnel. Initially, data for stationary airfoil held fixed at various angles-of-attack are gathered. Flow visualization pictures as well as detailed cross-sectional properties areobtained at various streamwise locations using hot-wire anemometry. Data include mean velocity, streamwise vorticity as well as various turbulent stresses. Preliminary data are also acquired for periodically pitched airfoil. These results are briefly presented in this extended abstract.
2-D Flow Numerical Solution for Airfoil and Hovercraft in Ground Effect
1978-12-01
REPRODUCE LEGIBLYo 2-D FLOW NUMERICAL SOLUTION FOR AIRFOIL AND HOVERCRAFT i GROUND EFECT THESIS AFIT/GAE/AA/78D-6 Itzhak Dvir Maj IAF $Approved for public...release; distribution unlimited. ,Li i i -AFIT/GAE/AA/78D-6 ’-~’ 2-D FLOW NUMERICAL SOLUTION FOR AIRFOIL AND HOVERCRAFT IN GROUND EFFECTS i .. THESIS ...Flow Numerical Solution for Airfoil and M.S. Thesis Hovercraft in Ground Effect I .C PERFORMING ORG. RCPOIFT NUM=BER 7. AUTHOR(&) S. CONTRAC’ Or GR
Determination of instantaneous pressure in a transonic base flow using four-pulse tomographic PIV
Blinde, P.L.; Lynch, K.P.; Schrijer, F.F.J.; Van Oudheusden, B.W.
2015-01-01
A tomographic four-pulse PIV system is used in a transonic axisymmetric base flow experiment at a nominal free stream Mach number of 0.7, with the objective to obtain flow acceleration and pressure data. The PIV system, consisting of two double-pulse lasers and twelve cameras, allows acquiring two
Numerical solution of inviscid transonic flow through 3D axial blade row
Energy Technology Data Exchange (ETDEWEB)
Fort, J.; Fuerst, J.; Halama, J.; Kozel, K. [CTU Prague (Czech Republic). Dept. of Technical Mathematics
2000-07-01
Presented paper deals with numerical solution of 3D inviscid transonic flow through axial cascades. Two different finite volume methods are mentioned. Authors show a comparison of both methods using results computed for the stator and the rotor cascades. A role of inlet parameters and body forces in the case of a rotor flow has been also investigated. (orig.)
Directory of Open Access Journals (Sweden)
Cui Michael M.
2005-01-01
Full Text Available To reduce vibration and noise level, the impeller and diffuser blade numbers inside an industrial compressor are typically chosen without common divisors. The shapes of volutes or collectors in these compressors are also not axis-symmetric. When impeller blades pass these asymmetric structures, the flow field in the compressor is time-dependent and three-dimensional. To obtain a fundamental physical understanding of these three-dimensional unsteady flow fields and assess their impact on the compressor performance, the flow field inside the compressors needs to be studied as a whole to include asymmetric and unsteady interaction between the compressor components. In the current study, a unified three-dimensional numerical model was built for a transonic centrifugal compressor including impeller, diffusers, and volute. HFC 134a was used as the working fluid. The thermodynamic and transport properties of the refrigerant gas were modeled by the Martin-Hou equation of state and power laws, respectively. The three-dimensional unsteady flow field was simulated with a Navier-Stokes solver using the k−ϵ turbulent model. The overall performance parameters are obtained by integrating the field quantities. Both the unsteady flow field and the overall performance are analyzed comparatively for each component. The compressor was tested in a water chiller system instrumented to obtain both the overall performance data and local flow-field quantities. The experimental and numerical results agree well. The correlation between the overall compressor performance and local flow-field quantities is defined. The methodology developed and data obtained in these studies can be applied to the centrifugal compressor design and optimization.
Nonlinear power flow feedback control for improved stability and performance of airfoil sections
Wilson, David G.; Robinett, III, Rush D.
2013-09-03
A computer-implemented method of determining the pitch stability of an airfoil system, comprising using a computer to numerically integrate a differential equation of motion that includes terms describing PID controller action. In one model, the differential equation characterizes the time-dependent response of the airfoil's pitch angle, .alpha.. The computer model calculates limit-cycles of the model, which represent the stability boundaries of the airfoil system. Once the stability boundary is known, feedback control can be implemented, by using, for example, a PID controller to control a feedback actuator. The method allows the PID controller gain constants, K.sub.I, K.sub.p, and K.sub.d, to be optimized. This permits operation closer to the stability boundaries, while preventing the physical apparatus from unintentionally crossing the stability boundaries. Operating closer to the stability boundaries permits greater power efficiencies to be extracted from the airfoil system.
The computation of the post-stall behavior of a circulation controlled airfoil
Linton, Samuel W.
1993-01-01
The physics of the circulation controlled airfoil is complex and poorly understood, particularly with regards to jet stall, which is the eventual breakdown of lift augmentation by the jet at some sufficiently high blowing rate. The present paper describes the numerical simulation of stalled and unstalled flows over a two-dimensional circulation controlled airfoil using a fully implicit Navier-Stokes code, and the comparison with experimental results. Mach numbers of 0.3 and 0.5 and jet total to freestream pressure ratios of 1.4 and 1.8 are investigated. The Baldwin-Lomax and k-epsilon turbulence models are used, each modified to include the effect of strong streamline curvature. The numerical solutions of the post-stall circulation controlled airfoil show a highly regular unsteady periodic flowfield. This is the result of an alternation between adverse pressure gradient and shock induced separation of the boundary layer on the airfoil trailing edge.
Airfoil data sensitivity analysis for actuator disc simulations used in wind turbine applications
DEFF Research Database (Denmark)
Nilsson, Karl; Breton, Simon-Philippe; Sørensen, Jens Nørkær
2014-01-01
how the mean characteristics of wake flow, mean power production and thrust depend on the choice of airfoil data and blade geometry. In order to simulate realistic conditions, pre-generated turbulence and wind shear are imposed in the computational domain. Using three different turbulence intensities......To analyse the sensitivity of blade geometry and airfoil characteristics on the prediction of performance characteristics of wind farms, large-eddy simulations using an actuator disc (ACD) method are performed for three different blade/airfoil configurations. The aim of the study is to determine...... of airfoil data in ACD simulations is not crucial if the intention of the simulations is to compute mean wake characteristics using a turbulent inflow....
Impact of uncertainty in airfoil characteristics on wind turbine extreme loads
DEFF Research Database (Denmark)
Abdallah, Imad; Natarajan, Anand; Sørensen, John Dalsgaard
2015-01-01
Wind tunnel test measurements to characterize the static lift and drag coefficients of airfoils used in wind turbine blades are shown to possess large uncertainties, which leads to uncertainties in the aerodynamic loads on the rotor. In this paper a rational stochastic model is proposed to quantify...... the uncertainty in airfoil static lift and drag coefficients based on field and wind tunnel data, aero-servoelastic calculations and engineering judgment. The stochastic model is subsequently used to assess the effect of the uncertainty in airfoil static lift and drag coefficients on the prediction of extreme...... loads and structural reliability of large wind turbines. It is shown that the uncertainty in the static airfoil data has a significant impact on the prediction of extreme loads effects and structural reliability depending on the component, operating conditions (stand-still versus power production...
Aerodynamic Characteristics of SC1095 and SC1094 R8 Airfoils
National Research Council Canada - National Science Library
Bousman, William G
2003-01-01
.... Measurements of the section lift, drag, and pitching moment have been obtained in ten wind tunnel tests for the SC1095 airfoil, and in five of these tests, measurements have also been obtained for the SC1094 R8...
Directory of Open Access Journals (Sweden)
Řidký Václav
2014-03-01
Full Text Available The work is devoted to 3D and 2D parallel numerical computation of pressure and velocity fields around an elastically supported airfoil self-oscillating due to interaction with the airflow. Numerical solution is computed in the OpenFOAM package, an open-source software package based on finite volume method. Movement of airfoil is described by translation and rotation, identified from experimental data. A new boundary condition for the 2DOF motion of the airfoil was implemented. The results of numerical simulations (velocity are compared with data measured in a wind tunnel, where a physical model of NACA0015 airfoil was mounted and tuned to exhibit the flutter instability. The experimental results were obtained previously in the Institute of Thermomechanics by interferographic measurements in a subsonic wind tunnel in Nový Knín.
Numerical simulation of dimples in airfoil using MATLAB
Booma Devi, P.; Shah, Dilip A.
2017-05-01
The Aircraft wing is a point of important research which poses greater challenge in terms of aerodynamic efficiency. The flow separation control method is addressed in classical aerodynamics methods. This study focuses on influence of dimples on controlling the flow and also increasing the aerodynamic efficiency. The periodic process of placing the cavities on the wing starting from root to tip controls the flow separation. The linear variation of characteristic curve provides the information about the flow separation and control of flow on upper surface of the airfoil.These different shapes are utilized viz., Square, Rectangle and Triangle. The numerical simulation is carried out in using MATLAB package. Preliminary analysis on the flow separation is carried out focuses on laminar flow separation, which has the influence on the overall lift generation and drag generation.
Compliance effects on dynamically pitching wind turbine airfoils
Magstadt, Andrew S.
The effects of elastic compliance in dynamically pitching wind turbine blades have been investigated. A numerical model guided wind tunnel testing, which used unsteady surface pressure measurements and phase-locked Particle Imaging Velocimetry to gather aerodynamic information. Using a torsionally compliant member, aeroelastic effects on the unsteady aerodynamics were compared against the results from a corresponding rigidly pitching airfoil to isolate the effects of compliance. The novel experimental apparatus and data acquisition techniques developed at the University of Wyoming showed that the presence of compliance can alter flow-field structures and increase dynamic loading. The high sensitivity of this nonlinear system suggests the formation of fluid-structure instabilities in large-scale turbines and demonstrates the potential for aerodynamic control as a means to mitigate adverse loading effects and improve wind turbine efficiency.
International Nuclear Information System (INIS)
Bak, Christian
2007-01-01
In this paper the influence of different key parameters in aerodynamic wind turbine rotor design on the power efficiency, C p , and energy production has been investigated. The work was divided into an analysis of 2D airfoils/blade sections and of entire rotors. In the analysis of the 2D airfoils it was seen that there was a maximum of the local C p for airfoils with finite maximum C l /C d values. The local speed ratio should be between 2.4 and 3.8 for airfoils with maximum c l /c d between 50 and 200, respectively, to obtain maximum local C p . Also, the investigation showed that Re had a significant impact on CP and especially for Re p for rotors was made with three blades and showed that with the assumption of constant maximum c l /c d along the entire blade, the design tip speed ratio changed from X=6 to X=12 for c l /cd=50 and c l /c d =200, respectively, with corresponding values of maximum c p of 0.46 and 0.525. An analysis of existing rotors re-designed with new airfoils but maintaining the absolute thickness distribution to maintain the stiffness showed that big rotors are more aerodynamic efficient than small rotors caused by higher Re. It also showed that the design tip speed ratio was very dependent on the rotor size and on the assumptions of the airfoil flow being fully turbulent (contaminated airfoil) or free transitional (clean airfoil). The investigations showed that rotors with diameter D=1.75m, should be designed for X around 5.5, whereas rotors with diameter D=126m, should be designed for Xbetween 6.5 and 8.5, depending on the airfoil performance
The flow of an incompressible electroconductive fluid past a thin airfoil. The parabolic profile
Directory of Open Access Journals (Sweden)
Adrian CARABINEANU
2014-04-01
Full Text Available We study the two-dimensional steady flow of an ideal incompressible perfectly conducting fluid past an insulating thin parabolic airfoil. We consider the linearized Euler and Maxwell equations and Ohm's law. We use the integral representations for the velocity, magnetic induction and pressure and the boundary conditions to obtain an integral equation for the jump of the pressure across the airfoil. We give some graphic representations for the lift coefficient, velocity and magnetic induction.
Family of airfoil shapes for rotating blades. [for increased power efficiency and blade stability
Noonan, K. W. (Inventor)
1983-01-01
An airfoil which has particular application to the blade or blades of rotor aircraft such as helicopters and aircraft propellers is described. The airfoil thickness distribution and camber are shaped to maintain a near zero pitching moment coefficient over a wide range of lift coefficients and provide a zero pitching moment coefficient at section Mach numbers near 0.80 and to increase the drag divergence Mach number resulting in superior aircraft performance.
Toward large eddy simulation of turbulent flow over an airfoil
Choi, Haecheon
1993-01-01
The flow field over an airfoil contains several distinct flow characteristics, e.g. laminar, transitional, turbulent boundary layer flow, flow separation, unstable free shear layers, and a wake. This diversity of flow regimes taxes the presently available Reynolds averaged turbulence models. Such models are generally tuned to predict a particular flow regime, and adjustments are necessary for the prediction of a different flow regime. Similar difficulties are likely to emerge when the large eddy simulation technique is applied with the widely used Smagorinsky model. This model has not been successful in correctly representing different turbulent flow fields with a single universal constant and has an incorrect near-wall behavior. Germano et al. (1991) and Ghosal, Lund & Moin have developed a new subgrid-scale model, the dynamic model, which is very promising in alleviating many of the persistent inadequacies of the Smagorinsky model: the model coefficient is computed dynamically as the calculation progresses rather than input a priori. The model has been remarkably successful in prediction of several turbulent and transitional flows. We plan to simulate turbulent flow over a '2D' airfoil using the large eddy simulation technique. Our primary objective is to assess the performance of the newly developed dynamic subgrid-scale model for computation of complex flows about aircraft components and to compare the results with those obtained using the Reynolds average approach and experiments. The present computation represents the first application of large eddy simulation to a flow of aeronautical interest and a key demonstration of the capabilities of the large eddy simulation technique.
Active Flow Control and Global Stability Analysis of Separated Flow Over a NACA 0012 Airfoil
Munday, Phillip M.
The objective of this computational study is to examine and quantify the influence of fundamental flow control inputs in suppressing flow separation over a canonical airfoil. Most flow control studies to this date have relied on the development of actuator technology, and described the control input based on specific actuators. Taking advantage of a computational framework, we generalize the inputs to fundamental perturbations without restricting inputs to a particular actuator. Utilizing this viewpoint, generalized control inputs aim to aid in the quantification and support the design of separation control techniques. This study in particular independently introduces wall-normal momentum and angular momentum to the separated flow using swirling jets through model boundary conditions. The response of the flow field and the surface vorticity fluxes to various combinations of actuation inputs are examined in detail. By closely studying different variables, the influence of the wall-normal and angular momentum injections on separated flow is identified. As an example, open-loop control of fully separated, incompressible flow over a NACA 0012 airfoil at alpha = 6° and 9° with Re = 23,000 is examined with large-eddy simulations. For the shallow angle of attack alpha = 6°, the small recirculation region is primarily affected by wall-normal momentum injection. For a larger separation region at alpha = 9°, it is observed that the addition of angular momentum input to wall-normal momentum injection enhances the suppression of flow separation. Reducing the size of the separated flow region significantly impacts the forces, and in particular reduces drag and increases lift on the airfoil. It was found that the influence of flow control on the small recirculation region (alpha = 6°) can be sufficiently quantified with the traditional coefficient of momentum. At alpha = 9°, the effects of wall-normal and angular momentum inputs are captured by modifying the standard
Mechanism of Water Droplet Breakup Near the Leading Edge of an Airfoil
Vargas, Mario; Sor, Suthyvann; Magarino, Adelaida, Garcia
2012-01-01
This work presents results of an experimental study on droplet deformation and breakup near the leading edge of an airfoil. The experiment was conducted in the rotating rig test cell at the Instituto Nacional de Tecnica Aeroespacial (INTA) in Madrid, Spain. The airfoil model was placed at the end of the rotating arm and a monosize droplet generator produced droplets that fell from above, perpendicular to the path of the airfoil. The interaction between the droplets and the airfoil was captured with high speed imaging and allowed observation of droplet deformation and breakup as the droplet approached the airfoil near the stagnation line. Image processing software was used to measure the position of the droplet centroid, equivalent diameter, perimeter, area, and the major and minor axes of an ellipse superimposed over the deforming droplet. The horizontal and vertical displacement of each droplet against time was also measured, and the velocity, acceleration, Weber number, Bond number, Reynolds number, and the drag coefficients were calculated along the path of the droplet to the beginning of breakup. Droplet deformation is defined and studied against main parameters. The high speed imaging allowed observation of the actual mechanism of breakup and identification of the sequence of configurations from the initiation of the breakup to the disintegration of the droplet. Results and comparisons are presented for droplets of diameters in the range of 500 to 1800 microns, and airfoil velocities of 70 and 90 m/sec.
Drag Coefficient of Water Droplets Approaching the Leading Edge of an Airfoil
Vargas, Mario; Sor, Suthyvann; Magarino, Adelaida Garcia
2013-01-01
This work presents results of an experimental study on droplet deformation and breakup near the leading edge of an airfoil. The experiment was conducted in the rotating rig test cell at the Instituto Nacional de Tecnica Aeroespacial (INTA) in Madrid, Spain. An airfoil model was placed at the end of the rotating arm and a monosize droplet generator produced droplets that fell from above, perpendicular to the path of the airfoil. The interaction between the droplets and the airfoil was captured with high speed imaging and allowed observation of droplet deformation and breakup as the droplet approached the airfoil near the stagnation line. Image processing software was used to measure the position of the droplet centroid, equivalent diameter, perimeter, area, and the major and minor axes of an ellipse superimposed over the deforming droplet. The horizontal and vertical displacement of each droplet against time was also measured, and the velocity, acceleration, Weber number, Bond number, Reynolds number, and the drag coefficients were calculated along the path of the droplet to the beginning of breakup. Results are presented and discussed for drag coefficients of droplets with diameters in the range of 300 to 1800 micrometers, and airfoil velocities of 50, 70 and 90 meters/second. The effect of droplet oscillation on the drag coefficient is discussed.
Derivation of airfoil characteristics for the LM 19.1 blade based on 3D CFD rotor calculations
Energy Technology Data Exchange (ETDEWEB)
Bak, C.; Soerensen, N.N.; Madsen, H.A. [Risoe National Lab., Roskilde (Denmark)
1999-03-01
Airfoil characteristics for the LM 19.1 blade are derived from 3D CFD computations on a full-scale 41-m rotor. Based on 3D CFD the force distributions on the blades are determined, from which airfoil characteristics are derived using the momentum theory. The final airfoil characteristics are constructed using both wind tunnel measurements and 3D CFD. Compared to 2D wind tunnel measurements they show a low lift in stall for the airfoil sections at the tip. At the airfoil sections at the inner part of the blade, they show a high lift in stall. At about 60% radius the lift agrees well to 2D wind tunnel measurements. Aero-elastic calculations using the final airfoil characteristics show good agreement to measured power and flap moments. Furthermore, a fatigue load analysis shows a reduction of up to 15% of the load compared to commonly used data. (au)
WIND TUNNEL RESEARCH ON THE INFLUENCE OF ACTIVE AIRFLOW ON THE LIFT FORCE GENERATED BY THE AIRFOIL
Directory of Open Access Journals (Sweden)
Paweł Magryta
2013-09-01
Full Text Available The paper discusses the results of wind tunnel tests of airfoils with additional active airflow applied to their upper surfaces. These studies were carried out for a range of velocities up to 28 m/s in an open wind tunnel. Several types of airfoils selected for the examination feature different geometries and are widely applied in today’s aviation industry. The changes in the lift and drag force generated by these airfoils were recorded during the study. The test bench for the tests was equipped with a compressor and a vacuum pump to enable airflow through some holes on the airfoil upper surface. A rapid prototyping method and a 3D printer based on a powder printing technique were applied to print the airfoils. All of their surfaces were subject to surface grinding to smooth their external surfaces. The wind tunnel tests with and without active airflow applied to airfoils are summarised in the paper.
Erickson, Gary E.; Schreiner, John A.; Rogers, Lawrence W.
1989-01-01
Slender wing vortex flows at subsonic, transonic, and supersonic speeds were investigated in a 6 x 6 ft wind tunnel. Test data obtained include off-body and surface flow visualizations, wing upper surface static pressure distributions, and six-component forces and moments. The results reveal the transition from the low-speed classical vortex regime to the transonic regime, beginning at a freestream Mach number of 0.60, where vortices coexist with shock waves. It is shown that the onset of core breakdown and the progression of core breakdown with the angle of attack were sensitive to the Mach number, and that the shock effects at transonic speeds were reduced by the interaction of the wing and the lead-edge extension (LEX) vortices. The vortex strengths and direct interaction of the wing and LEX cores (cores wrapping around each other) were found to diminish at transonic and supersonic speeds.
Rivers, Melissa; Quest, Juergen; Rudnik, Ralf
2015-01-01
Experimental aerodynamic investigations of the NASA Common Research Model have been conducted in the NASA Langley National Transonic Facility, the NASA Ames 11-ft wind tunnel, and the European Transonic Wind Tunnel. In the NASA Ames 11-ft wind tunnel, data have been obtained at only a chord Reynolds number of 5 million for a wing/body/tail = 0 degree incidence configuration. Data have been obtained at chord Reynolds numbers of 5, 19.8 and 30 million for the same configuration in the National Transonic Facility and in the European Transonic Facility. Force and moment, surface pressure, wing bending and twist, and surface flow visualization data were obtained in all three facilities but only the force and moment and surface pressure data are presented herein.
A Design of Experiments Investigation of Offset Streams for Supersonic Jet Noise Reduction
Henderson, Brenda; Papamoschou, Dimitri
2014-01-01
An experimental investigation into the noise characteristics of a dual-stream jet with four airfoils inserted in the fan nozzle was conducted. The intent of the airfoils was to deflect the fan stream relative to the core stream and, therefore, impact the development of the secondary potential core and noise radiated in the peak jet-noise direction. The experiments used a full-factorial Design of Experiments (DoE) approach to identify parameters and parameter interactions impacting noise radiation at two azimuthal microphone array locations, one of which represented a sideline viewing angle. The parameters studied included airfoil angle-of-attack, airfoil azimuthal location within the fan nozzle, and airfoil axial location relative to the fan-nozzle trailing edge. Jet conditions included subsonic and supersonic fan-stream Mach numbers. Heated jets conditions were simulated with a mixture of helium and air to replicate the exhaust velocity and density of the hot jets. The introduction of the airfoils was shown to impact noise radiated at polar angles in peak-jet noise direction and to have no impact on noise radiated at small and broadside polar angles and to have no impact on broadband-shock-associated noise. The DoE analysis showed the main effects impacting noise radiation at sideline-azimuthal-viewing angles included airfoil azimuthal angle for the airfoils on the lower side of the jet near the sideline array and airfoil trailing edge distance (with airfoils located at the nozzle trailing edge produced the lowest sound pressure levels). For an array located directly beneath the jet (and on the side of the jet from which the fan stream was deflected), the main effects impacting noise radiation included airfoil angle-of-attack and airfoil azimuthal angle for the airfoils located on the observation side of the jet as well and trailing edge distance. Interaction terms between multiple configuration parameters were shown to have significant impact on the radiated
Addy, Harold E., Jr.; Broeren, Andy P.; Potapczuk, Mark G.; Lee, Sam; Guffond, Didier; Montreuil, Emmanuel; Moens, Frederic
2016-01-01
in the IRT. From these molds, castings were made that closely replicated the features of the accreted ice. The castings were then mounted on the full-scale model in the F1 tunnel, and aerodynamic performance measurements were made using model surface pressure taps, the facility force balance system, and a large wake rake designed specifically for these tests. Tests were run over a range of Reynolds and Mach numbers. For each run, the model was rotated over a range of angles-of-attack that included airfoil stall. The benchmark data collected during these campaigns were, and continue to be, used for various purposes. The full-scale data form a unique, ice-accretion and associated aerodynamic performance dataset that can be used as a reference when addressing concerns regarding the use of subscale ice-accretion data to assess full-scale icing effects. Further, the data may be used in the development or enhancement of both ice-accretion prediction codes and computational fluid dynamic codes when applied to study the effects of icing. Finally, as was done in the wider study, the data may be used to help determine the level of geometric fidelity needed for artificial ice used to assess aerodynamic degradation due to aircraft icing. The structured, multifaceted approach used in this research effort provides a unique perspective on the aerodynamic effects of aircraft icing. The data presented in this report are available in electronic form upon formal approval by proper NASA and ONERA authorities.
Characteristics of merging shear layers and turbulent wakes of a multi-element airfoil
Adair, Desmond; Horne, W. Clifton
1988-02-01
Flow characteristics in the vicinity of the trailing edge of a single-slotted airfoil flap are presented and analyzed. The experimental arrangement consisted of a NACA 4412 airfoil equipped with a NACA 4415 flap whose angle of deflection was 21.8 deg. The flow remained attached over the model surfaces except in the vicinity of the flap trailing edge where a small region of boundary-layer separation extended over the aft 7 percent of flap chord. The flow was complicated by the presence of a strong, initially inviscid jet emanating from the slot between airfoil and flap, and a gradual merging of the main airfoil wake and flap suction-side boundary layer. Downstream of the flap, the airfoil and flap wakes fully merged to form an asymmetrical curved wake. The airfoil configuration was tested at an angle of attack of 8.2 deg, at a Mach number of 0.09, and a chord based Reynolds number of 1.8 x 10 to the 6th power in the Ames Research Center 7- by 10-Foot Wind Tunnel. Surface pressure measurements were made on the airfoil and flap and on the wind tunnel roof and floor. It was estimated that the wall interference increased the C sub L by 7 percent and decreased the C sub M by 4.5 percent. Velocity characteristics were quantified using hot-wire anemometry in regions of flow with preferred direction and low turbulence intensity. A 3-D laser velocimeter was used in regions of flow recirculation and relatively high turbulence intensity.
Addition of Passive Dynamics to a Flapping Airfoil to Improve Performance
Asselin, Daniel; Young, Jay; Williamson, C. H. K.
2017-11-01
Animals which fly or swim typically employ flapping motions of their wings and fins in order to produce thrust and to maneuver. Small, unmanned vehicles might also exploit such motions and are of considerable interest for the purposes of surveillance, environmental monitoring, and search and rescue. Flapping refers to a combination of pitch and heave and has been shown to provide good thrust and efficiency (Read, et al. 2003) when both axes are independently controlled (an Active-Active system). In this study, we examine the performance of an airfoil actuated only in the heave direction but allowed to pitch passively under the control of a torsion spring (an Active-Passive system). The presence of the spring is simulated in software using a force-feedback control system called Cyber-Physical Fluid Dynamics, or CPFD (Mackowski & Williamson 2011, 2015, 2016). Adding passive pitch to active heave provides significantly improved thrust and efficiency compared with heaving alone, especially when the torsion spring stiffness is selected so that the system operates near resonance (in an Active-Passive system). In many cases, values of thrust and efficiency are comparable to or better than those obtained with two actively controlled degrees of freedom. By using carefully-designed passive dynamics in the pitch direction, we can eliminate one of the two actuators, saving cost, complexity, and weight, while maintaining performance. This work was supported by the Air Force Office of Scientific Research Grant No. FA9550-15-1-0243, monitored by Dr. Douglas Smith.
Predicting the aeroelastic behavior of a wind-tunnel model using transonic small disturbance theory
Silva, Walter A.; Bennett, Robert M.
1990-01-01
The CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) code, developed at the NASA-Langley Research Center, is applied to the Active Flexible Wing (AFW) wind-tunnel model for prediction of the model's transonic aeroelastic behavior. Static aeroelastic solutions using CAP-TSD are computed. Dynamic (flutter) analyses are then performed as perturbations about the static aeroelastic deformations of the AFW. The accuracy of the static aeroelastic procedure is investigated by comparing analytical results to those from AFW wind-tunnel experiments. Dynamic results are presented in the form of root loci at different Mach numbers for a heavy gas and for air test mediums. The resultant flutter boundaries for both gases, and the effects of viscous damping and angle of attack on the flutter boundary in air, are also presented.
Aerodynamics of a highly irregular body at transonic speeds-Analysis of STRATOS flight data.
Guerster, Markus; Walter, Ulrich
2017-01-01
In this paper, we analyze the trajectory and body attitude data of Felix Baumgartner's supersonic free fall through the atmosphere on October 14, 2012. As one of us (UW) was scientific advisor to the Red Bull Stratos team, the analysis is based on true body data (body mass, wetted pressure suit surface area) and actual atmospheric data from weather balloon measurements. We also present a fully developed theoretical analysis and solution of atmospheric free fall. By matching the flight data against this solution, we are able to derive and track the drag coefficient CD from the subsonic to the transonic and supersonic regime, and back again. Although the subsonic drag coefficient is the expected CD = 0.60 ± 0.05, surprisingly the transonic compressibility drag coefficient is only 19% of the expected value. We provide a plausible explanation for this unexpected result.
Numerical solution of Euler and Navier-Stokes equations for 2D transonic problems
Hulek, T.; Hunek, M.; Kozel, K.
1992-12-01
The present contribution is a numerical solution of Euler and Navier-Stokes equations for 2D transonic flow problems using several different numerical methods. The time marching cell centered and cell vertex finite volume methods were used for both flow models. Various explicit multistage Runge-Kutta methods (RK methods) were applied for inviscid flows and these methods were also used for numerical solution of incompressible and compressinble Navier-Stokes equations.
LES Investigation of Wake Development in a Transonic Fan Stage for Aeroacoustic Analysis
Hah, Chunill; Romeo, Michael
2017-01-01
Detailed development of the rotor wake and its interaction with the stator are investigated with a large eddy simulation (LES). Typical steady and unsteady Navier-Stokes approaches (RANS and URANS) do not calculate wake development accurately and do not provide all the necessary information for an aeroacoustic analysis. It is generally believed that higher fidelity analysis tools are required for an aeroacoustic investigation of transonic fan stages.
Directory of Open Access Journals (Sweden)
Nožička Jiří
2012-04-01
Full Text Available This paper is part of the development of an airfoil for an unmanned aerial vehicle (UAV with internal propulsion system; the investigation involves the analysis of the aerodynamic performance for the gliding condition of two-dimensional airfoil models which have been tested. This development is based on the modification of a selected airfoil from the NACA four digits family. The modification of this base airfoil was made in order to create a blowing outlet with the shape of a step on the suction surface since the UAV will have an internal propulsion system. This analysis involved obtaining the lift, drag and pitching moment coefficients experimentally for the situation where there is not flow through the blowing outlet, called the no blowing condition by means of wind tunnel tests. The methodology to obtain the forces experimentally was through an aerodynamic wire balance. Obtained results were compared with numerical results by means of computational fluid dynamics (CFD from references and found in very good agreement. Finally, a selection of the airfoil with the best aerodynamic performance is done and proposed for further analysis including the blowing condition.
Velazquez, Luis; Nožička, Jiří; Vavřín, Jan
2012-04-01
This paper is part of the development of an airfoil for an unmanned aerial vehicle (UAV) with internal propulsion system; the investigation involves the analysis of the aerodynamic performance for the gliding condition of two-dimensional airfoil models which have been tested. This development is based on the modification of a selected airfoil from the NACA four digits family. The modification of this base airfoil was made in order to create a blowing outlet with the shape of a step on the suction surface since the UAV will have an internal propulsion system. This analysis involved obtaining the lift, drag and pitching moment coefficients experimentally for the situation where there is not flow through the blowing outlet, called the no blowing condition by means of wind tunnel tests. The methodology to obtain the forces experimentally was through an aerodynamic wire balance. Obtained results were compared with numerical results by means of computational fluid dynamics (CFD) from references and found in very good agreement. Finally, a selection of the airfoil with the best aerodynamic performance is done and proposed for further analysis including the blowing condition.
2D CFD Analysis of an Airfoil with Active Continuous Trailing Edge Flap
Jaksich, Dylan; Shen, Jinwei
2014-11-01
Efficient and quieter helicopter rotors can be achieved through on-blade control devices, such as active Continuous Trailing-Edge Flaps driven by embedded piezoelectric material. This project aims to develop a CFD simulation tool to predict the aerodynamic characteristics of an airfoil with CTEF using open source code: OpenFOAM. Airfoil meshes used by OpenFOAM are obtained with MATLAB scripts. Once created it is possible to rotate the airfoil to various angles of attack. When the airfoil is properly set up various OpenFOAM properties, such as kinematic viscosity and flow velocity, are altered to achieve the desired testing conditions. Upon completion of a simulation, the program gives the lift, drag, and moment coefficients as well as the pressure and velocity around the airfoil. The simulation is then repeated across multiple angles of attack to give full lift and drag curves. The results are then compared to previous test data and other CFD predictions. This research will lead to further work involving quasi-steady 2D simulations incorporating NASTRAN to model aeroelastic deformation and eventually to 3D aeroelastic simulations. NSF ECE Grant #1358991 supported the first author as an REU student.
Experimental study of ice accretion effects on aerodynamic performance of an NACA 23012 airfoil
Directory of Open Access Journals (Sweden)
Sohrab Gholamhosein Pouryoussefi
2016-06-01
Full Text Available In this paper, the effects of icing on an NACA 23012 airfoil have been studied. Experiments were applied on the clean airfoil, runback ice, horn ice, and spanwise ridge ice at a Reynolds number of 0.6 × 106 over angles of attack from −8° to 20°, and then results are compared. Generally, it is found that ice accretion on the airfoil can contribute to formation of a flow separation bubble on the upper surface downstream from the leading edge. In addition, it is made clear that spanwise ridge ice provides the greatest negative effect on the aerodynamic performance of the airfoil. In this case, the stall angle drops about 10° and the maximum lift coefficient reduces about 50% which is hazardous for an airplane. While horn ice leads to a stall angle drop of about 4° and a maximum lift coefficient reduction to 21%, runback ice has the least effect on the flow pattern around the airfoil and the aerodynamic coefficients so as the stall angle decreases 2° and the maximum lift reduces about 8%.
Modular turbine airfoil and platform assembly with independent root teeth
Campbell, Christian X; Davies, Daniel O; Eng, Darryl
2013-07-30
A turbine airfoil (22E-H) extends from a shank (23E-H). A platform (30E-H) brackets or surrounds a first portion of the shank (23E-H). Opposed teeth (33, 35) extend laterally from the platform (30E-H) to engage respective slots (50) in a disk. Opposed teeth (25, 27) extend laterally from a second portion of the shank (29) that extends below the platform (30E-H) to engage other slots (52) in the disk. Thus the platform (30E-H) and the shank (23E-H) independently support their own centrifugal loads via their respective teeth. The platform may be formed in two portions (32E-H, 34E-H), that are bonded to each other at matching end-walls (37) and/or via pins (36G) passing through the shank (23E-H). Coolant channels (41, 43) may pass through the shank beside the pins (36G).
Force Measurement Improvements to the National Transonic Facility Sidewall Model Support System
Goodliff, Scott L.; Balakrishna, Sundareswara; Butler, David; Cagle, C. Mark; Chan, David; Jones, Gregory S.; Milholen, William E., II
2016-01-01
The National Transonic Facility is a transonic pressurized cryogenic facility. The development of the high Reynolds number semi-span capability has advanced over the years to include transonic active flow control and powered testing using the sidewall model support system. While this system can be used in total temperatures down to -250Â F for conventional unpowered configurations, it is limited to temperatures above -60Â F when used with powered models that require the use of the high-pressure air delivery system. Thermal instabilities and non-repeatable mechanical arrangements revealed several data quality shortfalls by the force and moment measurement system. Recent modifications to the balance cavity recirculation system have improved the temperature stability of the balance and metric model-to-balance hardware. Changes to the mechanical assembly of the high-pressure air delivery system, particularly hardware that interfaces directly with the model and balance, have improved the repeatability of the force and moment measurement system. Drag comparisons with the high-pressure air system removed will also be presented in this paper.
Blended-Wing-Body Transonic Aerodynamics: Summary of Ground Tests and Sample Results
Carter, Melissa B.; Vicroy, Dan D.; Patel, Dharmendra
2009-01-01
The Blended-Wing-Body (BWB) concept has shown substantial performance benefits over conventional aircraft configuration with part of the benefit being derived from the absence of a conventional empennage arrangement. The configuration instead relies upon a bank of trailing edge devices to provide control authority and augment stability. To determine the aerodynamic characteristics of the aircraft, several wind tunnel tests were conducted with a 2% model of Boeing's BWB-450-1L configuration. The tests were conducted in the NASA Langley Research Center's National Transonic Facility and the Arnold Engineering Development Center s 16-Foot Transonic Tunnel. Characteristics of the configuration and the effectiveness of the elevons, drag rudders and winglet rudders were measured at various angles of attack, yaw angles, and Mach numbers (subsonic to transonic speeds). The data from these tests will be used to develop a high fidelity simulation model for flight dynamics analysis and also serve as a reference for CFD comparisons. This paper provides an overview of the wind tunnel tests and examines the effects of Reynolds number, Mach number, pitch-pause versus continuous sweep data acquisition and compares the data from the two wind tunnels.
Wu, Chung-Hua
1993-01-01
This report represents a general theory applicable to axial, radial, and mixed flow turbomachines operating at subsonic and supersonic speeds with a finite number of blades of finite thickness. References reflect the evolution of computational methods used, from the inception of the theory in the 50's to the high-speed computer era of the 90's. Two kinds of relative stream surfaces, S(sub 1) and S(sub 2), are introduced for the purpose of obtaining a three-dimensional flow solution through the combination of two-dimensional flow solutions. Nonorthogonal curvilinear coordinates are used for the governing equations. Methods of computing transonic flow along S(sub 1) and S(sub 2) stream surfaces are given for special cases as well as for fully three-dimensional transonic flows. Procedures pertaining to the direct solutions and inverse solutions are presented. Information on shock wave locations and shapes needed for computations are discussed. Experimental data from a Deutsche Forschungs- und Versuchsanstalt fur Luft- und Raumfahrt e.V. (DFVLR) rotor and from a Chinese Academy of Sciences (CAS) transonic compressor rotor are compared with the computed flow properties.
Lorenzoni, V.; Violato, D.; Scarano, F.
2010-01-01
Time-resolved Tomographic PIV was used to characterize the flow around the leading edge of a NACA 0012 airfoil in rod-airfoil configuration at ReD = 3500. The volumetric approach at relatively high temporal resolution allows the measurement of the evolution of the 3D vortical structures constituting
DEFF Research Database (Denmark)
Yilmaz, Özlem Ceyhan; Pires, Oscar; Munduate, Xabier
2017-01-01
This paper summarizes the results of a blind test campaign organized in the AVATAR project to predict the high Reynolds number performance of a wind turbine airfoil for wind turbine applications. The DU00-W-210 airfoil was tested in the DNW-HDG pressurized wind tunnel in order to investigate the ...
Wind tunnel test on airfoil Riso-B1-18 with an Active Trailing Edge Flap
DEFF Research Database (Denmark)
Bak, Christian; Gaunaa, Mac; Andersen, Peter Bjørn
2010-01-01
A wind tunnel test of the wind turbine airfoil Risø-B1-18 equipped with an Active Trailing Edge Flap (ATEF) was carried out. The ATEF was 9% of the total chord, made of piezo electric actuators attached to the trailing edge of a non-deformable airfoil and actuated using an (electric) amplifier...
Aerodynamic Design of a Tailless Aeroplan
Directory of Open Access Journals (Sweden)
J. Friedl
2001-01-01
Full Text Available The paper presents an aerodynamic analysis of a one-seat ultralight (UL tailless aeroplane named L2k, with a very complicated layout. In the first part, an autostable airfoil with a low moment coefficient was chosen as a base for this problem. This airfoil was refined and modified to satisfy the design requirements. The computed aerodynamic characteristics of the airfoils for different Reynolds numbers (Re were compared with available experimental data. XFOIL code was used to perform the computations. In the second part, a computation of wing characteristics was carried out. All calculated cases were chosen as points on the manoeuvring and gust envelope. The vortex lattice method was used with consideration of fuselage and winglets for very complicated wing geometry. The PMW computer program developed at IAE was used to perform the computations. The computed results were subsequently used for structural and strength analysis and design.
Riley, Donald R.
2015-01-01
This paper contains a collection of some results of four individual studies presenting calculated numerical values for airfoil aerodynamic stability derivatives in unseparated inviscid incompressible flow due separately to angle-of-attack, pitch rate, flap deflection, and airfoil camber using a discrete vortex method. Both steady conditions and oscillatory motion were considered. Variables include the number of vortices representing the airfoil, the pitch axis / moment center chordwise location, flap chord to airfoil chord ratio, and circular or parabolic arc camber. Comparisons with some experimental and other theoretical information are included. The calculated aerodynamic numerical results obtained using a limited number of vortices provided in each study compared favorably with thin airfoil theory predictions. Of particular interest are those aerodynamic results calculated herein (such as induced drag) that are not readily available elsewhere.
Directory of Open Access Journals (Sweden)
M.A. Neshat
2015-12-01
Full Text Available In this paper, the simultaneous effects of the sweep and lean of the blades in one stage of a transonic compressor on its performance have been investigated. Then, with the help of numerical solution, fluid flows over these two modified geometries generated from the original sample were analyzed. Considering the applied constraints, the two generated rotor geometries have different geometrical characteristics; so that in rotor No. 1, the blade has a backward sweep and it is less affected by lean, while in the modified rotor No. 2, the blade has a forward sweep and it is more affected by lean. In the first sample, it is observed that the stage efficiency increases by 0.5% for operating design, while the stall margin reduces, and the chocking mass flow rate diminishes by 1.5%. Also regarding the second modified blade, the results indicate that the stall margin increases, the choking flow rate at the nominal rotational speed of the stage increases by 0.18% and the stage efficiency increases by 1%. The comparison of numerical results also shows that, in the first modified rotor, the pressure ratio of the stage diminishes by 0.01%; while in the second sample, the pressure ratio of the stage increases by the same amount. These results were then compared with the experimental results, showing a good agreement.
Goldman, Louis J.
1993-01-01
An advanced laser anemometer (LA) was used to measure the axial and tangential velocity components in an annular cascade of turbine stator vanes operating at transonic flow conditions. The vanes tested were based on a previous redesign of the first-stage stator in a two-stage turbine for a high-bypass-ratio engine. The vanes produced 75 deg of flow turning. Tests were conducted on a 0.771-scale model of the engine-sized stator. The advanced LA fringe system employed an extremely small 50-micron diameter probe volume. Window correction optics were used to ensure that the laser beams did not uncross in passing through the curved optical access port. Experimental LA measurements of velocity and turbulence were obtained at the mean radius upstream of, within, and downstream of the stator vane row at an exit critical velocity ratio of 1.050 at the hub. Static pressures were also measured on the vane surface. The measurements are compared, where possible, with calculations from a three-dimensional inviscid flow analysis. Comparisons were also made with the results obtained previously when these same vanes were tested at the design exit critical velocity ratio of 0.896 at the hub. The data are presented in both graphical and tabulated form so that they can be readily compared against other turbomachinery computations.
CFD simulation of flow-induced vibration of an elastically supported airfoil
Directory of Open Access Journals (Sweden)
Šidlof Petr
2016-01-01
Full Text Available Flow-induced vibration of lifting or control surfaces in aircraft may lead to catastrophic consequences. Under certain circumstances, the interaction between the airflow and the elastic structure may lead to instability with energy transferred from the airflow to the structure and with exponentially increasing amplitudes of the structure. In the current work, a CFD simulation of an elastically supported NACA0015 airfoil with two degrees of freedom (pitch and plunge coupled with 2D incompressible airflow is presented. The geometry of the airfoil, mass, moment of inertia, location of the centroid, linear and torsional stiffness was matched to properties of a physical airfoil model used for wind-tunnel measurements. The simulations were run within the OpenFOAM computational package. The results of the CFD simulations were compared with the experimental data.
Wind-Tunnel Investigation of an NACA 23012 Airfoil with Various Arrangements of Slotted Flaps
Wenzinger, Carl J; Harris , Thomas A
1939-01-01
An investigation was made in the 7 by 10-foot wind tunnel and in the variable-density wind tunnel of the NACA 23012 airfoil with various slotted-flap arrangements. The purpose of the investigation in the 7 by 10-foot wind tunnel was to determine the airfoil section aerodynamic characteristics as affected by flap shape, slot shape, and flap location. The flap position for maximum lift; polars for arrangements favorable for take-off and climb; and complete lift, drag, and pitching-moment characteristics for selected optimum arrangements were determined. The best arrangements were tested in the variable-density tunnel at an effective Reynolds number of 8,000,000. In addition, data from both wind tunnels are included for plain, split, external-airfoil, and Fowler flaps for purposes of comparison.
Turbine airfoil cooling system with cooling systems using high and low pressure cooling fluids
Energy Technology Data Exchange (ETDEWEB)
Marsh, Jan H.; Messmann, Stephen John; Scribner, Carmen Andrew
2017-10-25
A turbine airfoil cooling system including a low pressure cooling system and a high pressure cooling system for a turbine airfoil of a gas turbine engine is disclosed. In at least one embodiment, the low pressure cooling system may be an ambient air cooling system, and the high pressure cooling system may be a compressor bleed air cooling system. In at least one embodiment, the compressor bleed air cooling system in communication with a high pressure subsystem that may be a snubber cooling system positioned within a snubber. A delivery system including a movable air supply tube may be used to separate the low and high pressure cooling subsystems. The delivery system may enable high pressure cooling air to be passed to the snubber cooling system separate from low pressure cooling fluid supplied by the low pressure cooling system to other portions of the turbine airfoil cooling system.
Predicting the Extreme Loads on a Wind Turbine Considering Uncertainty in Airfoil Data
DEFF Research Database (Denmark)
Abdallah, Imad; Natarajan, Anand; Sørensen, John Dalsgaard
2014-01-01
The sources contributing to uncertainty in a wind turbine blade static airfoil data include wind tunnel testing, CFD calculations, 3D rotational corrections based on CFD or emprircal models, surface roughness corrections, Reynolds number corrections, expansion to the full 360-degree angle of attack...... range, validation by full scale measurements, and geometric distortions of the blade during manufacturing and under loading. In this paper a stochastic model of the static airfoil data is proposed to supplement the prediction of extreme loads effects for large wind turbines. It is shown...... that the uncertainty in airfoil data can have e significant impact on the prediction of extreme loads effects depending on the component, and the correlation along the span of the blade....
Hybrid immersed boundary method for airfoils with a trailing-edge flap
DEFF Research Database (Denmark)
Zhu, Wei Jun; Behrens, Tim; Shen, Wen Zhong
2013-01-01
In this paper, a hybrid immersed boundary technique has been developed for simulating turbulent flows past airfoils with moving trailing-edge flaps. Over the main fixed part of the airfoil, the equations are solved using a standard body-fitted finite volume technique, whereas the moving trailing......-edge flap is simulated using the immersed boundary method on a curvilinear mesh. An existing in-house-developed flow solver is employed to solve the incompressible Reynolds-Averaged Navier-Stokes equations together with the k-ω turbulence model. To achieve consistent wall boundary conditions at the immersed...... boundaries the k-ωturbulence model is modified and adapted to the local conditions associated with the immersed boundary method. The obtained results show that the hybrid approach is an efficient and accurate method for solving turbulent flows past airfoils with a trailing-edge flap and that flow control...
Comparisons between LES and wind tunnel hot-wire measurements of a NACA 0015 airfoil
DEFF Research Database (Denmark)
Zhu, Wei Jun; Shen, Wen Zhong; Bertagnolio, Franck
2012-01-01
is constructed in a wind tunnel similar as the condition where the experiments were carried out. The goal of this study is to validate the LES model against detailed measurements. The simulations are performed with in-house EllipSys3D code on high performance computers. Numerical study are focused...... on the stability and accuracy of the LES simulations on various mesh configurations. The spanwise grid spacing was found important to produce correct flow disturbance along the airfoil span, which further affects the turbulent energy distribution.......Large-eddy simulations (LES) are carried out for flow over a NACA 0015 airfoil at AoA = 8o and chord based Reynolds number of 1.71106. To accurately simulate the complex flow at the suction side of the airfoil, a reasonably large number of grid points are required. The computational mesh...
Numerical simulation of the divergence of a wind turbine airfoil : part 2
Energy Technology Data Exchange (ETDEWEB)
Ramdenee, D.; Minea, I.S.; Tardiff d' Hamonville, T.; Illinca, A. [Quebec Univ., Rimouski, PQ (Canada). Laboratoire de Recherche en Energie Eolienne
2010-07-01
The development of larger, more flexible wind turbine blades is creating the need for an improved understanding of the mechanisms surrounding unsteady flow-structure interactions. This 2-part study used computational fluid dynamics (CFD) to identify and model the aerodynamic and aeroelastic phenomena around wind turbine blades. Aeroelastic divergence was modelled using coupled aerodynamic and elastic models with an ANSYS software program. The fluid-structure interactions of an NACA0012 airfoil were simulated in order to determine the divergence phenomenon created by aerodynamic loads and transient fluid flow. The airfoil profile was fixed and exempted from all rotational degrees of liberty while being subjected to a constant flow of velocity. The fixing was then removed and the constant flow was compared with a shock wave on the airfoil profile. The profile then oscillated with damped amplitude due to the aerodynamic damping imposed. Results of the analysis will be compared with results obtained in future studies. 7 refs., 2 tabs., 6 figs.
Chitta, Varun
Lifting surfaces of unmanned aerial vehicles (UAV) are often operated in low Reynolds number (Re) ranges, wherein the transition of boundary layer from laminar-to-turbulent plays a more significant role than in high-Re aerodynamics applications. This poses a challenge for traditional computational fluid dynamics (CFD) simulations, since typical modeling approaches assume either fully laminar or fully turbulent flow. In particular, the boundary layer state must be accurately predicted to successfully determine the separation behavior which significantly influences the aerodynamic characteristics of the airfoil. Reynolds-averaged Navier-Stokes (RANS) based CFD simulations of an elliptic airfoil are performed for time-varying angles of attack, and results are used to elucidate relevant flow physics and aerodynamic data for an elliptic airfoil under realistic operating conditions. Results are also used to evaluate the performance of several different RANS-based turbulence modeling approaches for this class of flowfield.
An experimental study of an airfoil with a bio-inspired leading edge device at high angles of attack
Mandadzhiev, Boris A.; Lynch, Michael K.; Chamorro, Leonardo P.; Wissa, Aimy A.
2017-09-01
Robust and predictable aerodynamic performance of unmanned aerial vehicles at the limits of their design envelope is critical for safety and mission adaptability. Deployable aerodynamic surfaces from the wing leading or trailing edges are often used to extend the aerodynamic envelope (e.g. slats and flaps). Birds have also evolved feathers at the leading edge (LE) of their wings, known as the alula, which enables them to perform high angles of attack maneuvers. In this study, a series of wind tunnel experiments are performed to quantify the effect of various deployment parameters of an alula-like LE device on the aerodynamic performance of a cambered airfoil (S1223) at stall and post stall conditions. The alula relative angle of attack, measured from the mean chord of the airfoil, is varied to modulate tip-vortex strength, while the alula deflection angle is varied to modulate the distance between the tip vortex and the wing surface. Integrated lift force measurements were collected at various alula-inspired device configurations. The effect of the alula-inspired device on the boundary layer velocity profile and turbulence intensity were investigated through hot-wire anemometer measurements. Results show that as alula deflection angle increases, the lift coefficient also increase especially at lower alula relative angles of attack. Moreover, at post stall wing angles of attack, the wake velocity deficit is reduced in the presence of alula device, confirming the mitigation of the wing adverse pressure gradient. The results are in strong agreement with measurements taken on bird wings showing delayed flow reversal and extended range of operational angles of attack. An engineered alula-inspired device has the potential to improve mission adaptability in small unmanned air vehicles during low Reynolds number flight.
Effectiveness of spoilers on the GA(W)-1 airfoil with a high performance Fowler flap
Wentz, W. H., Jr.
1975-01-01
Two-dimensional wind-tunnel tests were conducted to determine effectiveness of spoilers applied to the GA(W)-1 airfoil. Tests of several spoiler configurations show adequate control effectiveness with flap nested. It is found that providing a vent path allowing lower surface air to escape to the upper surface as the spoiler opens alleviates control reversal and hysteresis tendencies. Spoiler cross-sectional shape variations generally have a modest influence on control characteristics. A series of comparative tests of vortex generators applied to the (GA-W)-1 airfoil show that triangular planform vortex generators are superior to square planform vortex generators of the same span.
Turbine airfoil with an internal cooling system having vortex forming turbulators
Lee, Ching-Pang
2014-12-30
A turbine airfoil usable in a turbine engine and having at least one cooling system is disclosed. At least a portion of the cooling system may include one or more cooling channels having a plurality of turbulators protruding from an inner surface and positioned generally nonorthogonal and nonparallel to a longitudinal axis of the airfoil cooling channel. The configuration of turbulators may create a higher internal convective cooling potential for the blade cooling passage, thereby generating a high rate of internal convective heat transfer and attendant improvement in overall cooling performance. This translates into a reduction in cooling fluid demand and better turbine performance.
Aerodynamic behaviour of NREL S826 airfoil at Re=100,000
DEFF Research Database (Denmark)
Chivaee, Hamid Sarlak; Mikkelsen, Robert Flemming; Sarmast, Sasan
2014-01-01
at Fluid Mechanics laboratory of the Technical University of Denmark (DTU). Lift coefficient is obtained from the forge gauge measurements while the drag is measured according to the integration of the wake profiles downstream of the airfoil. The pressure distribution is measured by a set of pressure taps......, there is a better agreement between the drag measurements and computations. It is concluded that LES computations are able to capture the lift and drag polars as well as the pressure distribution around the airfoil with an acceptable accuracy....
Computational Analysis of the 2415-3S Airfoil Aerodynamic Performance
Directory of Open Access Journals (Sweden)
Luis Velázquez-Araque
2014-02-01
Full Text Available This paper deals with the numerical simulation of the two-dimensional, incompressible, steady air flow past an airfoil for a solar powered unmanned aerial vehicle (UAV with internal propulsion system. This airfoil results from a NACA 2415 four digits family base airfoil modification [7] and has a propulsive outlet with the shape of a step on the suction surface. The analysis involved the airfoil's aerodynamic performance which meant obtaining lift, drag and pitching moment coefficient curves as a function of the angle of attack (AOA for the condition where the engine of the UAV is turned off called the gliding condition and also for the blowing propulsive condition by means computational fluid dynamics. The computational domain has been discretised using a structured mesh of 188 x 200 tetrahedral elements. The RNG k-Ε model is utilized to describe the turbulent flow process as it was followed in [5]. The simulations were held at a Reynolds number of 300000. Results allowed obtaining lift and drag forces and pitching moment coefficient and also the location of the separation and reattachment points in some cases by means of the wall shear stress on the suction surface as well as velocity contours and streamlines for both conditions at different angles of attack, from 0 to 16 degrees with the smallest increment of 4 degrees. Finally, results from both cases were compared and the influence of the propulsive flow on the aerodynamic characteristics of the airfoil has been analysed turning out that it improves significantly the performance of the airfoil reaching values up to 1,8 times in terms of lift at high angles of attack. [5] Rhie C.M., Chow W.L., Numerical Study of the Turbulent Flow Past an Airfoil with Trailing Edge Separation, AIAA Journal, Vol. 21, No. 11, 1983. [7] Velazquez L., Nožička J, Kulhanek R., Oil and Smoke Flow Visualization past Two-Dimensional Airfoils for an Unmanned Aerial Vehicle, in The 11th Asian Symposium of
An Iterative Method for Estimating Airfoil Deformation due to Solid Particle Erosion
Directory of Open Access Journals (Sweden)
Valeriu DRAGAN
2014-04-01
Full Text Available Helicopter blades are currently constructed with composite materials enveloping honeycomb cores with only the leading and trailing edges made of metal alloys. In some cases, the erosive wear of the bound between the composite skin and metallic leading edge leads to full blade failure. It is therefore the goal of this paper to provide a method for simulating the way an airfoil is deformed through the erosion process. The method involves computational fluid dynamics simulations, scripts for automatic meshing and spreadsheet calculators for estimating the erosion and, ultimately, the airfoil deformation. Further work could include more complex meshing scripts allowing the use of similar methods for turbo-machineries.
Nonlinear angle control of a sectioned airfoil by using shape memory alloys
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Abreu G.
2014-01-01
Full Text Available The present work illustrates an application of shape memory alloys and nonlinear controller applied to the active angular control of a sectioned airfoil. The main objective of the proposed control system is to modify the shape of the profile based on a reference angle. The change of the sectioned airfoil angle is resultant by the effect of shape memory of the alloy due to heating of the wire caused by an electric current that changes its temperature by Joule effect. Considering the presence of plant’s nonlinear effects, especially in the mathematical model of the alloy, this work proposes the application of an on-off control system.
Aerodynamics and Optimal Design of Biplane Wind Turbine Blades
Chiu, Phillip
In order to improve energy capture and reduce the cost of wind energy, in the past few decades wind turbines have grown significantly larger. As their blades get longer, the design of the inboard region (near the blade root) becomes a trade-off between competing structural and aerodynamic requirements. State-of-the-art blades require thick airfoils near the root to efficiently support large loads inboard, but those thick airfoils have inherently poor aerodynamic performance. New designs are required to circumvent this design compromise. One such design is the "biplane blade", in which the thick airfoils in the inboard region are replaced with thinner airfoils in a biplane configuration. This design was shown previously to have significantly increased structural performance over conventional blades. In addition, the biplane airfoils can provide increased lift and aerodynamic efficiency compared to thick monoplane inboard airfoils, indicating a potential for increased power extraction. This work investigates the fundamental aerodynamic aspects, aerodynamic design and performance, and optimal structural design of the biplane blade. First, the two-dimensional aerodynamics of biplanes with relatively thick airfoils are investigated, showing unique phenomena which arise as a result of airfoil thickness. Next, the aerodynamic design of the full biplane blade is considered. Two biplane blades are designed for optimal aerodynamic loading, and their aerodynamic performance quantified. Considering blades with practical chord distributions and including the drag of the mid-blade joint, it is shown that biplane blades have comparable power output to conventional monoplane designs. The results of this analysis also show that the biplane blades can be designed with significantly less chord than conventional designs, a characteristic which enables larger blade designs. The aerodynamic loads on the biplane blades are shown to be increased in gust conditions and decreased under
An experimental study of airfoil instability tonal noise with trailing edge serrations
Chong, Tze Pei; Joseph, Phillip F.
2013-11-01
This paper presents an experimental study of the effect of trailing edge serrations on airfoil instability noise. Detailed aeroacoustic measurements are presented of the noise radiated by an NACA-0012 airfoil with trailing edge serrations in a low to moderate speed flow under acoustical free field conditions. The existence of a separated boundary layer near the trailing edge of the airfoil at an angle of attack of 4.2 degree has been experimentally identified by a surface mounted hot-film arrays technique. Hot-wire results have shown that the saw-tooth surface can trigger a bypass transition and prevent the boundary layer from becoming separated. Without the separated boundary layer to act as an amplifier for the incoming Tollmien-Schlichting waves, the intensity and spectral characteristic of the radiated tonal noise can be affected depending upon the serration geometry. Particle Imaging Velocimetry (PIV) measurements of the airfoil wakes for a straight and serrated trailing edge are also reported in this paper. These measurements show that localized normal-component velocity fluctuations that are present in a small region of the wake from the laminar airfoil become weakened once serrations are introduced. Owing to the above unique characteristics of the serrated trailing edges, we are able to further investigate the mechanisms of airfoil instability tonal noise with special emphasis on the assessment of the wake and non-wake based aeroacoustic feedback models. It has been shown that the instability tonal noise generated at an angle of attack below approximately one degree could involve several complex mechanisms. On the other hand, the non-wake based aeroacoustic feedback mechanism alone is sufficient to predict all discrete tone frequencies accurately when the airfoil is at a moderate angle of attack. Larger Δf, which is defined as (fn+1-fn). In other words, a larger margin of velocity increase is required in order to "shift" the fn and fn+1 across fs
Lowry, John G
1941-01-01
An investigation was made in the NACA 7- by 10-foot wind tunnel to determine the effect of slot-lip location on the aerodynamic section characteristics of an NACA 23012 airfoil with a 30-percent-chord slotted flap. Tests were made with slot lips located at 90 and 100 percent of the airfoil chord and with two different flap shapes. The results are compared with a slotted flap previously developed by the National advisory Committee for Aeronautics with a slot lip located at 83 percent of the airfoil chord. The extension of the slot lip to the rear increased the section lift and pitching-moment coefficients. Comparisons made on a basis of pitching moment for a given tail length show that the Fowler type flap, lip extended to trailing edge of the airfoil, has the greatest section lift coefficient. For moderate tail lengths, 2 to 3 chord lengths, there was only a slight difference between the previously developed slotted flap and the slotted flap with slot lip extended to 90 percent of the airfoil chord. Of the three flaps tested, the Fowler flap had the lowest drag coefficient at high lift coefficients. The extension of the lower surface at the leading edge of the slot had a negligible effect on the profile drag of the airfoil-flap arrangement with the flap deflected when the lip terminated at 90 percent of the airfoil chord.
Vargas, Mario; Feo, Alex
2011-01-01
This work presents the results of an experimental study on droplet deformation and breakup near the leading edge of an airfoil. The experiment was conducted in the rotating rig test cell at the Instituto Nacional de Tecnica Aeroespacial (INTA) in Madrid, Spain. An airfoil model placed at the end of the rotating arm was moved at speeds of 50 to 90 m/sec. A monosize droplet generator was employed to produce droplets that were allowed to fall from above, perpendicular to the path of the airfoil at a given location. High speed imaging was employed to observe the interaction between the droplets and the airfoil. The high speed imaging allowed observation of droplet deformation and breakup as the droplet approached the airfoil near the stagnation line. A tracking software program was used to measure from the high speed movies the horizontal and vertical displacement of the droplet against time. The velocity, acceleration, Weber number, Bond number, Reynolds number, and the drag coefficients were calculated along the path of a given droplet from beginning of deformation to breakup and/or hitting the airfoil. Results are presented for droplets with a diameter of 490 micrometers at airfoil speeds of 50, 60, 70, 80 and 90 m/sec
Lift Optimization Study of a Multi-Element Three-Segment Variable Camber Airfoil
Kaul, Upender K.; Nguyen, Nhan T.
2016-01-01
This paper reports a detailed computational high-lift study of the Variable Camber Continuous Trailing Edge Flap (VCCTEF) system carried out to explore the best VCCTEF designs, in conjunction with a leading edge flap called the Variable Camber Krueger (VCK), for take-off and landing. For this purpose, a three-segment variable camber airfoil employed as a performance adaptive aeroelastic wing shaping control effector for a NASA Generic Transport Model (GTM) in landing and take-off configurations is considered. The objective of the study is to define optimal high-lift VCCTEF settings and VCK settings/configurations. A total of 224 combinations of VCK settings/configurations and VCCTEF settings are considered for the inboard GTM wing, where the VCCTEFs are configured as a Fowler flap that forms a slot between the VCCTEF and the main wing. For the VCK settings of deflection angles of 55deg, 60deg and 65deg, 18, 19 and 19 vck configurations, respectively, were considered for each of the 4 different VCCTEF deflection settings. Different vck configurations were defined by varying the horizontal and vertical distance of the vck from the main wing. A computational investigation using a Reynolds-Averaged Navier-Stokes (RANS) solver was carried out to complement a wind-tunnel experimental study covering three of these configurations with the goal of identifying the most optimal high-lift configurations. Four most optimal high-lift configurations, corresponding to each of the VCK deflection settings, have been identified out of all the different configurations considered in this study yielding the highest lift performance.
Colera, Manuel; Pérez-Saborid, Miguel
2017-09-01
A finite differences scheme is proposed in this work to compute in the time domain the compressible, subsonic, unsteady flow past an aerodynamic airfoil using the linearized potential theory. It improves and extends the original method proposed in this journal by Hariharan, Ping and Scott [1] by considering: (i) a non-uniform mesh, (ii) an implicit time integration algorithm, (iii) a vectorized implementation and (iv) the coupled airfoil dynamics and fluid dynamic loads. First, we have formulated the method for cases in which the airfoil motion is given. The scheme has been tested on well known problems in unsteady aerodynamics -such as the response to a sudden change of the angle of attack and to a harmonic motion of the airfoil- and has been proved to be more accurate and efficient than other finite differences and vortex-lattice methods found in the literature. Secondly, we have coupled our method to the equations governing the airfoil dynamics in order to numerically solve problems where the airfoil motion is unknown a priori as happens, for example, in the cases of the flutter and the divergence of a typical section of a wing or of a flexible panel. Apparently, this is the first self-consistent and easy-to-implement numerical analysis in the time domain of the compressible, linearized coupled dynamics of the (generally flexible) airfoil-fluid system carried out in the literature. The results for the particular case of a rigid airfoil show excellent agreement with those reported by other authors, whereas those obtained for the case of a cantilevered flexible airfoil in compressible flow seem to be original or, at least, not well-known.
Time-Dependent Effects of Glaze Ice on the Aerodynamic Characteristics of an Airfoil
Directory of Open Access Journals (Sweden)
Narges Tabatabaei
2018-01-01
Full Text Available The main objective of this study is to estimate the dynamic loads acting over a glaze-iced airfoil. This work studies the performance of unsteady Reynolds-averaged Navier-Stokes (URANS simulations in predicting the oscillations over an iced airfoil. The structure and size of time-averaged vortices are compared to measurements. Furthermore, the accuracy of a two-equation eddy viscosity turbulence model, the shear stress transport (SST model, is investigated in the case of the dynamic load analysis over a glaze-iced airfoil. The computational fluid dynamic analysis was conducted to investigate the effect of critical ice accretions on a 0.610 m chord NACA 0011 airfoil. Leading edge glaze ice accretion was simulated with flat plates (spoiler-ice extending along the span of the blade. Aerodynamic performance coefficients and pressure profiles were calculated and validated for the Reynolds number of 1.83 × 106. Furthermore, turbulent separation bubbles were studied. The numerical results confirm both time-dependent phenomena observed in previous similar measurements: (1 low-frequency mode, with a Strouhal number Sth≈0,013–0.02, and (2 higher frequency mode with a Strouhal number StL≈0,059–0.69. The higher frequency motion has the same characteristics as the shedding mode and the lower frequency motion has the flapping mode characteristics.
Cross-Validation of Numerical and Experimental Studies of Transitional Airfoil Performance
DEFF Research Database (Denmark)
Frere, Ariane; Hillewaert, Koen; Sarlak, Hamid
2015-01-01
The aerodynamic performance characteristic of airfoils are the main input for estimating wind turbine blade loading as well as annual energy production of wind farms. For transitional flow regimes these data are difficult to obtain, both experimentally as well as numerically, due to the very high...
Computations of droplet impingement on airfoils in two-phase flow
International Nuclear Information System (INIS)
Kim, Sang Dug; Song, Dong Joo
2005-01-01
The aerodynamic effects of leading-edge accretion can raise important safety concerns since the formulation of ice causes severe degradation in aerodynamic performance as compared with the clean airfoil. The objective of this study is to develop a numerical simulation strategy for predicting the particle trajectory around an MS-0317 airfoil in the test section of the NASA Glenn Icing Research Tunnel and to investigate the impingement characteristics of droplets on the airfoil surface. In particular, predictions of the mean velocity and turbulence diffusion using turbulent flow solver and continuous random walk method were desired throughout this flow domain in order to investigate droplet dispersion. The collection efficiency distributions over the airfoil surface in simulations with different numbers of droplets, various integration time-steps and particle sizes were compared with experimental data. The large droplet impingement data indicated the trends in impingement characteristics with respect to particle size; the maximum collection efficiency located at the upper surface near the leading edge, and the maximum value and total collection efficiency were increased as the particle size was increased. The extent of the area impinged on by particles also increased with the increment of the particle size, which is similar as compared with experimental data
The effect of electrohydrodynamic force on the lift coefficient of a NACA 0015 airfoil
Yusof, Y.; Hossain, A.; Abdullah, A. H.; Nasir, Rizal M. E.; Hamid, A.; Muthmainnah, N.; N, M.
2017-11-01
Lift, the force component that is perpendicular to the line of flight, is generated when a small aircraft moves through the air. With the help of the sets of flaps and slats on its wing, the pilot controls his aircraft manoeuvring in the air. In this study, we preferred to cut the drawbacks of the flaps system by introducing the electrohydrodynamic actuator. Widely known as plasma actuator, it is able to improve the induced lift force as well as the efficiency of a small aircraft system. A dielectric-barrier-discharge actuator using a 6 kV AC power supply was developed and tested on a NACA 0015 airfoil using copper as the electrodes and kapton as its dielectric component. The experimental results showed that it was successful in presenting a positive effect of the plasma actuator on the lift coefficient of the airfoil at smaller angle of attack, where enhancements ranged between 0.7% and 1.8%. However, at a higher angle, the results were not as swayed as it was desired since the energy exerted by the plasma actuator on the lift performance of the airfoil was inadequate. Further tests are needed using higher rated voltage supply and other equipment to improve the capability of the actuator in refining the aerodynamic performance of the airfoil.
Unsteady two-dimensional potential-flow model for thin variable geometry airfoils
DEFF Research Database (Denmark)
Gaunaa, Mac
2010-01-01
on the airfoil. This time-lag term can be approximated using an indicial function approach, making the practical calculation of the aerodynamic response numerically very efficient by use of Duhamel superposition. Furthermore, the indicial function expressions for the time-lag terms are formulated...
Predicting extreme loads effects on wind turbines considering uncertainty in airfoil data
DEFF Research Database (Denmark)
Abdallah, Imad; Natarajan, Anand; Sørensen, John Dalsgaard
2013-01-01
The sources contributing to uncertainty in a wind turbine blade static airfoil data include wind tunnel testing, CFD calculations, 3D rotational corrections based on CFD or empirical models, surface roughness corrections, Reynolds number corrections, expansion to the full 360-degree angle of atta...... that the uncertainty in airfoil data can have a significant impact on the prediction of extreme loads effects depending on the component, and the correlation along the span of the blade.......The sources contributing to uncertainty in a wind turbine blade static airfoil data include wind tunnel testing, CFD calculations, 3D rotational corrections based on CFD or empirical models, surface roughness corrections, Reynolds number corrections, expansion to the full 360-degree angle of attack...... range, validation by full scale measurements, and geometric distortions of the blade during manufacturing and under loading. In this paper a stochastic model of the static airfoil data is proposed to supplement the prediction of extreme loads effects for large wind turbines. It is shown...
Active Control of Flow Separation Over an Airfoil Using Synthetic Jets
You, D.; Moin, P.
We perform large-eddy simulation of turbulent flow separation over an airfoil and evaluate the effectiveness of synthetic jets as a separation control technique.The flow configuration consists of flow over a NACA 0015 airfoil at Reynolds number of 896,000 based on the airfoil chord length and freestream velocity.A small slot across the entire span connected to a cavity inside the airfoil is employed to produce oscillatory synthetic jets.Detailed flow structures inside the synthetic-jet actuator and the synthetic jet/cross-flow interaction are simulated using an unstructured-grid finite-volume large-eddy simulation solver.Simulation results are compared with the experimental data of Gilarranz et al.(J.Fluids Eng.127, pp.377-387 (2005)), and qualitative and quantitative agreements are obtained for both uncontrolled and controlled cases.As in the experiment, the present large-eddy simulation confirms that synthetic-jet actuation effectively delays the onset of flow separation and causes a significant increase in the lift coefficient.Modification of the blade boundary layer due to oscillatory blowing and suction and its role in separation control is discussed.
Aerodynamic behaviour of NREL S826 airfoil at Re=100,000
Sarlak, H.; Mikkelsen, R.; Sarmast, S.; Sørensen, J. N.
2014-06-01
This paper presents wind tunnel measurements of the NREL S826 airfoil at Reynolds number Re = 100,000 for angles of attack in a range of -10° to 25° the corresponding Large Eddy Simulation (LES) for selected angles of attack. The measurements have been performed at the low speed wind tunnel located at Fluid Mechanics laboratory of the Technical University of Denmark (DTU). Lift coefficient is obtained from the forge gauge measurements while the drag is measured according to the integration of the wake profiles downstream of the airfoil. The pressure distribution is measured by a set of pressure taps on the airfoil surface. The lift and drag polars are obtained from the LES computations using DTU's inhouse CFD solver, EllipSys3D, and good agreement is found between the measurement and the simulations. At high angles of attack, the numerical computations tend to over-predict the lift coefficients, however, there is a better agreement between the drag measurements and computations. It is concluded that LES computations are able to capture the lift and drag polars as well as the pressure distribution around the airfoil with an acceptable accuracy.
Estimation of morphing airfoil shape and aerodynamic load using artificial hair sensors
Butler, Nathan S.; Su, Weihua; Thapa Magar, Kaman S.; Reich, Gregory W.
2016-04-01
An active area of research in adaptive structures focuses on the use of continuous wing shape changing methods as a means of replacing conventional discrete control surfaces and increasing aerodynamic efficiency. Although many shape-changing methods have been used since the beginning of heavier-than-air flight, the concept of performing camber actuation on a fully-deformable airfoil has not been widely applied. A fundamental problem of applying this concept to real-world scenarios is the fact that camber actuation is a continuous, time-dependent process. Therefore, if camber actuation is to be used in a closed-loop feedback system, one must be able to determine the instantaneous airfoil shape as well as the aerodynamic loads at all times. One approach is to utilize a new type of artificial hair sensors developed at the Air Force Research Laboratory to determine the flow conditions surrounding deformable airfoils. In this work, the hair sensor measurement data will be simulated by using the flow solver XFoil, with the assumption that perfect data with no noise can be collected from the hair sensor measurements. Such measurements will then be used in an artificial neural network based process to approximate the instantaneous airfoil camber shape, lift coefficient, and moment coefficient at a given angle of attack. Various aerodynamic and geometrical properties approximated from the artificial hair sensor and artificial neural network system will be compared with the results of XFoil in order to validate the approximation approach.
A comparative study on the flow over an airfoil using transitional turbulence models
DEFF Research Database (Denmark)
Lin, Mou; Sarlak Chivaee, Hamid
2016-01-01
This work addresses the simulation of the flow over NREL S826 airfoil under a relatively low Reynolds number (Re = 1 × 105 ) using the CFD solvers OpenFoam and ANSYS Fluent. The flow is simulated using two different transition models, γ − Reθ and k − kL − ω model, and the results are examined...
Studi Eksperimen dan Numerik Pengaruh Penambahan Vortex Generator pada Airfoil NASA LS-0417
Directory of Open Access Journals (Sweden)
Ulul Azmi
2017-03-01
Full Text Available Separasi boundary layer merupakan fenomena penting yang mempengaruhi performansi airfoil. Salah satu upaya untuk menunda atau menghilangkan separasi aliran adalah meningkatkan momentum fluida untuk melawan adverse pressure dan tegangan geser permukaan. Hal ini mengakibatkan separasi aliran akan tertunda lebih ke belakang. Upaya tersebut dapat dilakukan dengan penambahan turbulent generator pada upper surface airfoil. Vortex generator (VG merupakan salah satu jenis turbulent generator yang dapat mempercepat transisi dari laminar boundary layer menjadi turbulent boundary layer. Oleh karena itu, penelitian ini bertujuan untuk mengetahui pengaruh jarak penempatan dan ketinggian VG terhadap perkembangan turbulent boundary layer sehingga dapat meningkatkan performansi airfoil. Penelitian ini dilakukan dengan eksperimen dan numerik pada Re = 1.41x105 dengan angle of attack 16°. Benda uji yang digunakan adalah airfoil NASA LS-0417 dengan dan tanpa VG. Variasi jarak penempatan dan ketinggian VG yaitu x/c = 0.1; 0.2; 0.3; 0.4 (h = 1 mm; 3 mm; 5 mm. Hasil yang didapatkan adalah variasi vortex generator paling optimal adalah vortex generator dengan x/c = 0.3 dan h = 1 mm dimana Nilai CL/CD mengalami kenaikan sebesar 14.337%.
DEFF Research Database (Denmark)
Yang, Hua; Shen, Wen Zhong; Xu, Haoran
2013-01-01
some models before they can be used in a BEM code. In this article, the airfoil data for the MEXICO (Model EXperiments in Controlled cOnditions) rotor are extracted from CFD (Computational Fluid Dynamics) results. The azimuthally averaged velocity is used as the sectional velocity to define the angle...
URANS simulations of separated flow with stall cells over an NREL S826 airfoil
DEFF Research Database (Denmark)
Sarlak Chivaee, Hamid; Nishino, T.; Sørensen, Jens Nørkær
2016-01-01
A series of wind tunnel measurements and oil flow visualization was recently carried out at the Technical University of Denmark in order to investigate flow characteristics over a 14% thick NREL S826 airfoil at low Reynolds numbers. This paper aims at presenting numerical simulations of the same ...
Computational Investigations on the Effects of Gurney Flap on Airfoil Aerodynamics.
Jain, Shubham; Sitaram, Nekkanti; Krishnaswamy, Sriram
2015-01-01
The present study comprises steady state, two-dimensional computational investigations performed on NACA 0012 airfoil to analyze the effect of Gurney flap (GF) on airfoil aerodynamics using k-ε RNG turbulence model of FLUENT. Airfoil with GF is analyzed for six different heights from 0.5% to 4% of the chord length, seven positions from 0% to 20% of the chord length from the trailing edge, and seven mounting angles from 30° to 120° with the chord. Computed values of lift and drag coefficients with angle of attack are compared with experimental values and good agreement is found at low angles of attack. In addition static pressure distribution on the airfoil surface and pathlines and turbulence intensities near the trailing edge are present. From the computational investigation, it is recommended that Gurney flaps with a height of 1.5% chord be installed perpendicular to chord and as close to the trailing edge as possible to obtain maximum lift enhancement with minimum drag penalty.
Control of Pitching Airfoil Aerodynamics by Vorticity Flux Modification using Active Bleed
Kearney, John; Glezer, Ari
2014-11-01
Distributed active bleed driven by pressure differences across a pitching airfoil is used to regulate the vorticity flux over the airfoil's surface and thereby to control aerodynamic loads in wind tunnel experiments. The range of pitch angles is varied beyond the static stall margin of the 2-D VR-7 airfoil at reduced pitching rates up to k = 0.42. Bleed is regulated dynamically using piezoelectric louvers between the model's pressure side near the trailing edge and the suction surface near the leading edge. The time-dependent evolution of vorticity concentrations over the airfoil and in the wake during the pitch cycle is investigated using high-speed PIV and the aerodynamic forces and moments are measured using integrated load cells. The timing of the dynamic stall vorticity flux into the near wake and its effect on the flow field are analyzed in the presence and absence of bleed using proper orthogonal decomposition (POD). It is shown that bleed actuation alters the production, accumulation, and advection of vorticity concentrations near the surface with significant effects on the evolution, and, in particular, the timing of dynamic stall vortices. These changes are manifested by alteration of the lift hysteresis and improvement of pitch stability during the cycle, while maintaining cycle-averaged lift to within 5% of the base flow level with significant implications for improvement of the stability of flexible wings and rotor blades. This work is supported by the Rotorcraft Center (VLRCOE) at Georgia Tech.
Identification of stiffness and damping characteristics of axial air-foil bearings
Arora, Vikas; Arora, Vikas; van der Hoogt, Peter; Aarts, Ronald G.K.M.; de Boer, Andries
2011-01-01
Air-foil bearings (AFBs) are self acting hydrodynamic bearings made from sheet metal foils comprised of at least two layers. The innermost “top foil” layer traps a gas pressure film that supports a load while the layer or layers underneath provide an elastic foundation. AFBs are currently used in
DEFF Research Database (Denmark)
Døssing, Mads
Time series of pressure fluctuations has been obtained using high frequency microphones distributed over the surface of airfoils undergoing wind tunnel tests in the LM Windtunnel, owned by ’LM Glasfiber’, Denmark. The present report describes the dataanalysis, with special attention given to tran...
Characterization of the Effect of Wing Surface Instrumentation on UAV Airfoil Performance
Ratnayake, Nalin A.
2009-01-01
Recently proposed flight research at NASA Dryden Flight Research Center (DFRC) has prompted study into the aerodynamic effects of modifications made to the surfaces of laminar airfoils. The research is focused on the high-aspect ratio, laminar-flow type wings commonly found on UAVs and other aircraft with a high endurance requirement. A broad range of instrumentation possibilities, such as structural, pressure, and temperature sensing devices may require the alteration of the airfoil outer mold line as part of the installation process. This study attempts to characterize the effect of installing this additiona1 instrumentation on key airfoil performance factors, such as transition location, lift and drag curves, and stall point. In particular, the general case of an airfoil that is channeled in the spanwise direction is considered, and the impact on key performance characteristics is assessed. Particular attention is focused on exploring the limits of channel depth and low-Reynolds number on performance and stall characteristics. To quantify the effect of increased skin friction due to premature transition caused by protruding or recessed instrumentation, two simplified, conservative scenarios are used to consider two potential sources of diaturbance: A) that leading edge alterations would cause linearly expanding areas (triangles) of turbulent flow on both surfaces of the wing upstream of the natural transition point, and B) that a channel or bump on the upper surface would trip turbulent flow across the whole upper surface upstream of the natural transition point. A potentially more important consideration than the skin friction drag increment is the change in overall airfoil performance due to the installation of instrumentation along most of the wingspan. To quantify this effect, 2D CFD simulations of the flow over a representative mid-span airfoil section were conducted in order to assess the change in lift and drag curves for the airfoil in the presence of
Effects of Leading Edge Defect on the Aerodynamic and Flow Characteristics of an S809 Airfoil.
Wang, Yan; Zheng, Xiaojing; Hu, Ruifeng; Wang, Ping
Unexpected performance degradation occurs in wind turbine blades due to leading edge defect when suffering from continuous impacts with rain drops, hails, insects, or solid particles during its operation life. To assess this issue, this paper numerically investigates the steady and dynamic stall characteristics of an S809 airfoil with various leading edge defects. More leading edge defect sizes and much closer to practical parameters are investigated in the paper. Numerical computation is conducted using the SST k-ω turbulence model, and the method has been validated by comparison with existed published data. In order to ensure the calculation convergence, the residuals for the continuity equation are set to be less than 10-7 and 10-6 in steady state and dynamic stall cases. The simulations are conducted with the software ANSYS Fluent 13.0. It is found that the characteristics of aerodynamic coefficients and flow fields are sensitive to leading edge defect both in steady and dynamic conditions. For airfoils with the defect thickness of 6%tc, leading edge defect has a relative small influence on the aerodynamics of S809 airfoil. For other investigated defect thicknesses, leading edge defect has much greater influence on the flow field structures, pressure coefficients and aerodynamic characteristics of airfoil at relative small defect lengths. For example, the lift coefficients decrease and drag coefficients increase sharply after the appearance of leading edge defect. However, the aerodynamic characteristics could reach a constant value when the defect length is large enough. The flow field, pressure coefficient distribution and aerodynamic coefficients do not change a lot when the defect lengths reach to 0.5%c,1%c, 2%c and 3%c with defect thicknesses of 6%tc, 12%tc,18%tc and 25%tc, respectively. In addition, the results also show that the critical defect length/thickness ratio is 0.5, beyond which the aerodynamic characteristics nearly remain unchanged. In
Adaptive Wing/Aerofoil Design Optimisation Using MOEA Coupled to Uncertainty Design Method
Lee, D.S.; Periaux, J.; Gonzalez, L.F.; Onate, E.; Qin, N.
2011-01-01
The use of adaptive wing/aerofoil designs is being considered as promising techniques in aeronautic/aerospace since they can reduce aircraft emissions, improve aerodynamic performance of manned or unmanned aircraft. The paper investigates the robust design and optimisation for one type of adaptive techniques; Active Flow Control (AFC) bump at transonic flow conditions on a Natural Laminar Flow (NLF) aerofoil designed to increase aerodynamic efficiency (especially high lift to drag ratio). The...
Analysis and Multipoint Design of the TCA Concept
Krist, Steven E.; Bauer, Steven X. S.; Buning, Pieter G.
1999-01-01
The goal in this effort is to analyze the baseline TCA concept at transonic and supersonic cruise, then apply the natural flow wing design concept to obtain multipoint performance improvements. Analyses are conducted with OVERFLOW, a Navier-Stokes code for overset grids, using PEGSUS to compute the interpolations between the overset grids.
3D Flow Past Transonic Turbine Cascade SE 1050-Experiment and Numerical Simulations
Czech Academy of Sciences Publication Activity Database
Šimurda, David; Fürst, J.; Luxa, Martin
2013-01-01
Roč. 22, č. 4 (2013), s. 311-319 ISSN 1003-2169. [International Symposium on Experimental and Computational Aerothermodynamics of Internal Flow s : ISAIF /11./. Shenzhen, 06.05.2013-11.05.2013] R&D Projects: GA ČR(CZ) GAP101/10/1329 Institutional support: RVO:61388998 Keywords : blade cascade * vortex structures * transonic flow * CFD Subject RIV: BK - Fluid Dynamics Impact factor: 0.348, year: 2013 http://link.springer.com/article/10.1007%2Fs11630-013-0629-7
Determination of instantaneous pressure in a transonic base flow using four-pulse tomographic PIV
Blinde, P.L.; Lynch, K.P.; Schrijer, F.F.J.; Van Oudheusden, B.W.
2015-01-01
A tomographic four-pulse PIV system is used in a transonic axisymmetric base flow experiment at a nominal free stream Mach number of 0.7, with the objective to obtain flow acceleration and pressure data. The PIV system, consisting of two double-pulse lasers and twelve cameras, allows acquiring two velocity fields with time separations as small as 2.5 ?s. A performance assessment is carried out and provides a typical average error estimate below 0.025 U? (0.3 voxel). The ability to use these v...
Prediction of forces and moments on finned bodies at high angle of attack in transonic flow
Energy Technology Data Exchange (ETDEWEB)
Oberkampf, W. L.
1981-04-01
This report describes a theoretical method for the prediction of fin forces and moments on bodies at high angle of attack in subsonic and transonic flow. The body is assumed to be a circular cylinder with cruciform fins (or wings) of arbitrary planform. The body can have an arbitrary roll (or bank) angle, and each fin can have individual control deflection. The method combines a body vortex flow model and lifting surface theory to predict the normal force distribution over each fin surface. Extensive comparisons are made between theory and experiment for various planform fins. A description of the use of the computer program that implements the method is given.
Unsteady separated boundary layer in a transonic diffuser flow with self-excited oscillations
Hsieh, T.; Coakley, T. J.
1986-01-01
A numerical investigation of two-dimensional unsteady boundary layer in a transonic diffuser flow with self-excited oscillations and strong flow separation by solving the compressible, Reynolds-averaged, thin-layer Navier-Stokes equations with two-equations turbulence model is described. Three different meshes with constant streamwise mesh distribution and varying vertical mesh distribution were used. Results obtained indicate that a refinement of mesh studied here has minimal effect on the mean boundary layer flow but significantly increases the amplitude of oscillation of all flow variables. Comparisons of unsteady wall pressure, velocity profile, terminal shock, and separation pocket among computations and with experiment are presented.
Numerical calculations of 2D transonic flow in GAMM channel and over the profile
Directory of Open Access Journals (Sweden)
Slouka Martin
2017-01-01
Full Text Available The aim of this work is to make a 2D numerical model of the solution of the transonic inviscid and viscous compressible flow around the profile. In a case of viscous flow several turbulent models are used. For the verification of the calculation Baldwin-Lomax model is compared with Wilcox k-omega model and SST turbulent model. Calculations are done in GAMM channel computational domain with 10% DCA profile and in the turbine cascade computational domain with 8% DCA profile. Numerical methods are based on a finite volume solution. Comparisons are done with the experimental data for the 8% DCA profile.
Numerical calculations of 2D transonic flow in GAMM channel and over the profile
Slouka, Martin; Kozel, Karel
The aim of this work is to make a 2D numerical model of the solution of the transonic inviscid and viscous compressible flow around the profile. In a case of viscous flow several turbulent models are used. For the verification of the calculation Baldwin-Lomax model is compared with Wilcox k-omega model and SST turbulent model. Calculations are done in GAMM channel computational domain with 10% DCA profile and in the turbine cascade computational domain with 8% DCA profile. Numerical methods are based on a finite volume solution. Comparisons are done with the experimental data for the 8% DCA profile.
Estimation of morphing airfoil shapes and aerodynamic loads using artificial hair sensors
Butler, Nathan Scott
An active area of research in adaptive structures focuses on the use of continuous wing shape changing methods as a means of replacing conventional discrete control surfaces and increasing aerodynamic efficiency. Although many shape-changing methods have been used since the beginning of heavier-than-air flight, the concept of performing camber actuation on a fully-deformable airfoil has not been widely applied. A fundamental problem of applying this concept to real-world scenarios is the fact that camber actuation is a continuous, time-dependent process. Therefore, if camber actuation is to be used in a closed-loop feedback system, one must be able to determine the instantaneous airfoil shape, as well as the aerodynamic loads, in real time. One approach is to utilize a new type of artificial hair sensors (AHS) developed at the Air Force Research Laboratory (AFRL) to determine the flow conditions surrounding deformable airfoils. In this study, AHS measurement data will be simulated by using the flow solver XFoil, with the assumption that perfect data with no noise can be collected from the AHS measurements. Such measurements will then be used in an artificial neural network (ANN) based process to approximate the instantaneous airfoil camber shape, lift coefficient, and moment coefficient at a given angle of attack. Additionally, an aerodynamic formulation based on the finite-state inflow theory has been developed to calculate the aerodynamic loads on thin airfoils with arbitrary camber deformations. Various aerodynamic properties approximated from the AHS/ANN system will be compared with the results of the finite-state inflow aerodynamic formulation in order to validate the approximation approach.
Aktharuzzaman, Md; Sarker, Md. Samad; Safa, Wasiul; Sharah, Nahreen; Salam, Md. Abdus
2017-12-01
Magnus effect is a phenomenon where pressure difference is created according to Bernoulli's effect due to induced velocity changes caused by a rotating object in a fluid. Using this concept, the idea of delaying boundary layer separation on airfoil by providing moving surface boundary layer control has been developed. In order to analyze the influence of Magnus effect on the aerodynamic performance of an airfoil, there is no alternative of developing an experimental setup. This paper aims to develop such an experimental setup which will be capable of analyzing the influence of Magnus effect on both symmetric and asymmetric airfoils by placing a cylinder at the leading edge. To provide arrangements for a rotating cylinder at the leading edge of airfoil, necessary modifications and additions have been done in the test section of an AF100 subsonic wind tunnel.
Vortex-induced vibrations of a DU96-W-180 airfoil at 90° angle of attack
DEFF Research Database (Denmark)
Skrzypinski, Witold Robert; Gaunaa, Mac; Sørensen, Niels N.
2014-01-01
This work presents an analysis of vortex-induced vibrations of a DU96-W-180 airfoil in deep stall at a 90 degrees angle of attack, based on 2D and 3D Reynolds Averaged Navier Stokes and 3D Detached Eddy Simulation unsteady Computational Fluid Dynamics computations with non-moving, prescribed motion...... and elastically mounted airfoil suspensions. Stationary vortex-shedding frequencies computed in 2D and 3D Computational Fluid Dynamics differed. In the prescribed motion computations, the airfoil oscillated in the direction of the chord line. Negative aerodynamic damping, found in both 2D and 3D Computational...... Fluid Dynamics computations with moving airfoil, showed in the vicinity of the stationary vortex-shedding frequency computed by 2D Computational Fluid Dynamics. A shorter time series was sufficient to verify the sign of the aerodynamic damping in the case of the elastic computations than the prescribed...
National Research Council Canada - National Science Library
Elmstrom, Michael
2004-01-01
A computational fluid dynamic (CFD) investigation is presented that provides predictions of the aerodynamic impact of uniform and non-uniform coatings applied to the leading edge of a compressor airfoil in a cascade. Using a NACA 65(12...
High-Reynolds Number Circulation Control Testing in the National Transonic Facility
Milholen, William E., II; Jones, Gregory S.; Chan, David T.; Goodliff, Scott L.
2012-01-01
A new capability to test active flow control concepts and propulsion simulations at high Reynolds numbers in the National Transonic Facility at the NASA Langley Research Center is being developed. The first active flow control experiment was completed using the new FAST-MAC semi-span model to study Reynolds number scaling effects for several circulation control concepts. Testing was conducted over a wide range of Mach numbers, up to chord Reynolds numbers of 30 million. The model was equipped with four onboard flow control valves allowing independent control of the circulation control plenums, which were directed over a 15% chord simple-hinged flap. Preliminary analysis of the uncorrected lift data showed that the circulation control increased the low-speed maximum lift coefficient by 33%. At transonic speeds, the circulation control was capable of positively altering the shockwave pattern on the upper wing surface and reducing flow separation. Furthermore, application of the technique to only the outboard portion of the wing demonstrated the feasibility of a pneumatic based roll control capability.
The evolution of whole field optical diagnostics for external transonic testing
Fry, K. A.; Bryanston-Cross, P.
1992-09-01
The diagnostic use of quantitative laser flow visualization techniques has increased rapidly over recent years. The limitations imposed by conventional single point techniques such as laser Doppler anemometry are addressed and how they have been overcome by the development of a new family of whole field measurement techniques is demonstrated. In particular near instantaneous whole field velocity data was obtained in a relatively hostile, industrial 2.74 m x 2.44 m transonic wind tunnel (TWT) at the Aircraft Research Association (ARA). The techniques were evaluated for their suitability for making quantitative measurements in the wing/pylon region of a model wing and engine combination. Three optical diagnostic techniques were successfully developed within the context of the ARA facility. The first technique, laser light sheet (LLS), combines the operation of a pulse laser and video capture system to provide a 'real time' visualization of the flow, whereas a second pulse laser technique, Particle Image Velocimetry (PIV) can be used to make specific quantitative whole field instantaneous velocity measurements. The third method, holography, was used to produce a stored three dimensional visualization of the unsteady and shock wave features of the transonic flow in the gully region. A description is made of their installation and operation, and examples are presented of current test results.
Energy Technology Data Exchange (ETDEWEB)
Delhaye, D.; Paniagua, G. [von Karman Institute for Fluid Dynamics, Turbomachinery and Propulsion Department, Rhode-Saint-Genese (Belgium); Fernandez Oro, J.M. [Universidad de Oviedo, Area de Mecanica de Fluidos, Gijon (Spain); Denos, R. [European Commission, Directorate General for Research, Brussels (Belgium)
2011-01-15
The paper presents the development and application of a three-sensor wedge probe to measure unsteady aerodynamics in a transonic turbine. CFD has been used to perform a detailed uncertainty analysis related to probe-induced perturbations, in particular the separation zones appearing on the wedge apex. The effects of the Reynolds and Mach numbers are studied using both experimental data together with CFD simulations. The angular range of the probe and linearity of the calibration maps are enhanced with a novel zonal calibration technique, used for the first time in compressible flows. The data reduction methodology is explained and demonstrated with measurements performed in a single-stage high-pressure turbine mounted in the compression tube facility of the von Karman Institute. The turbine was operated at subsonic and transonic pressure ratios (2.4 and 5.1) for a Reynolds number of 10{sup 6}, representative of modern engine conditions. Complete maps of the unsteady flow angle and rotor outlet Mach number are documented. These data allow the study of secondary flows and rotor trailing edge shocks. (orig.)
Usry, J. W.; Wallace, J. W.
1971-01-01
The forebody drag of a supercritical body of revolution was measured in free flight over a Mach number range of 0.85 to 1.05 and a Reynolds number range of 11.5 x 10 to the 6th power to 19.4 x 10 to the 6th power and was compared with wind-tunnel data. The forebody drag coefficient for a Mach number less than 0.96 was 0.111 compared with the wind-tunnel value of 0.103. A gradual increase in the drag occurred in the Langley 8-foot transonic pressure tunnel at a lower Mach number than in the Langley 16-foot transonic tunnel or in the free-flight test. The sharp drag rise occurred near Mach 0.98 in free flight whereas the rise occurred near Mach 0.99 in the Langley 16-foot transonic tunnel. The sharp rise was not as pronounced in the Langley 8-foot transonic pressure tunnel and was probably affected by tunnel-wall-interference effects. The increase occurred more slowly and at a higher Mach number. These results indicate that the drag measurements made in the wind tunnels near Mach 1 were significantly affected by the relative size of the model and the wind tunnel.
International Nuclear Information System (INIS)
Habbal, S.R.; Tsinganos, K.
1983-01-01
A new class of shock transitions are shown to arise in the transonic solutions of the steady isothermal solar wind equations when, for example, momentum deposition gives rise to multiple critical points in the flow. These shock transitions between critical solutions occur for a certain range of the parameters characterizing the momentum deposition function. In the presence of such shock transitions, the isothermal wind equations admit multiple transonic solutions, namely a continuous solution passing through an inner critical point and solutions involving a shock transition between critical solutions. These multiple transonic solutions have the same flow speed at the base but different supersonic flow speeds at infinity. An additional interesting feature of the isothermal equations is the equivalence of nonradial flow tube divergence and momentum addition in giving rise to multiple critical points and hence to multiple transonic solutions with shock transitions. The physical relevance of these properties for astrophysical systems such as the inner solar wind, flows in extragalactic jets and accretion discs are discussed
Mcghee, R. J.; Beasley, W. D.
1976-01-01
Wind tunnel tests were conducted to determine the effects of airfoil thickness-ratio on the low speed aerodynamic characteristics of an initial family of airfoils. The results were compared with theoretical predictions obtained from a subsonic viscous method. The tests were conducted over a Mach number range from 0.10 to 0.28. Chord Reynolds numbers varied from about 2.0 x 1 million to 9.0 x 1 million.
1980-03-31
VARIABLES IN FORCE COEFFICIENT PROGRAM A = matrix coefficient (Equation (C4)) ALFA = airfoil angle of attack (in degrees) ALFAC = airfoil angle of attack...8217,XOALFA’,5X,’ ALFAC ,96X,*CL’,8XOCLC7X’CD,X’CDC 17X, REL’ ,SX, ’ II s’ 5X, LB) ,3X, ’RVEL(F/S)’I XKA = -. 17598 XKB = -.14485 WAVE = 0.0 DO 10 11
Energy Technology Data Exchange (ETDEWEB)
Djouimaa, S. [Physical Department, Sciences Faculty, Batna University, Ave Chahid Boukhlouf Med, El-Hadi, 05000 Batna (Algeria); Messaoudi, L. [Mechanical Department, Engineering Sciences Faculty, Batna University, Ave Chahid Boukhlouf Med, El-Hadi, 05000 Batna (Algeria); Giel, Paul W. [QSS Group, Inc., NASAGlenn Research Center Cleveland, OH 44135 (United States)
2007-03-15
The objective of this study is to simulate the transonic gas turbine blade-to-blade compressible fluid flow. We are interested mainly in the determination of the pressure distribution around the blade. The particular blade architecture makes these simulations more complex due to the variety of phenomena induced by this flow. Our study is based on the experiment performed by Giel and colleagues. Tests were conducted in a linear cascade at the NASA Glenn Research Center. The test article was a turbine rotor with design flow turning of 136{sup o} and an axial chord of 12.7cm. Simulations were performed on an irregular quadratic structured grid with the FLUENT software package which solves the Navier-Stokes equations by using finite volume methods. Two-dimensional stationary numerical simulations were made under turbulent conditions allowing us to compare the characteristic flow effects of Reflecting Boundary Conditions (RBC) and Non-Reflecting Boundary Conditions (NRBC) newly implemented in FLUENT 6.0. Many simulations were made to compare different turbulence models: a one equation model (Spalart-Allmaras), several two-equation models (k-{epsilon}, RNG k-{epsilon}, Realizable k-{epsilon}, SST k-{omega}), and a Reynolds-stress model (RSM). Also examined were the effects of the inlet turbulence intensities (0.25% and 7%), the exit Mach numbers (1.0 and 1.3) and the inlet Reynolds numbers (0.5x10{sup 6} and 1x10{sup 6}). The results obtained show a good correlation with the experiment. (author)
Flegel, Ashlie B.; Welch, Gerard E.; Giel, Paul W.; Ames, Forrest E.; Long, Jonathon A.
2015-01-01
Two independent experimental studies were conducted in linear cascades on a scaled, two-dimensional mid-span section of a representative Variable Speed Power Turbine (VSPT) blade. The purpose of these studies was to assess the aerodynamic performance of the VSPT blade over large Reynolds number and incidence angle ranges. The influence of inlet turbulence intensity was also investigated. The tests were carried out in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility and at the University of North Dakota (UND) High Speed Compressible Flow Wind Tunnel Facility. A large database was developed by acquiring total pressure and exit angle surveys and blade loading data for ten incidence angles ranging from +15.8deg to -51.0deg. Data were acquired over six flow conditions with exit isentropic Reynolds number ranging from 0.05×106 to 2.12×106 and at exit Mach numbers of 0.72 (design) and 0.35. Flow conditions were examined within the respective facility constraints. The survey data were integrated to determine average exit total-pressure and flow angle. UND also acquired blade surface heat transfer data at two flow conditions across the entire incidence angle range aimed at quantifying transitional flow behavior on the blade. Comparisons of the aerodynamic datasets were made for three "match point" conditions. The blade loading data at the match point conditions show good agreement between the facilities. This report shows comparisons of other data and highlights the unique contributions of the two facilities. The datasets are being used to advance understanding of the aerodynamic challenges associated with maintaining efficient power turbine operation over a wide shaft-speed range.
Aerodynamic Performance of a NREL S809 Airfoil in an Air-Sand Particle Two-Phase Flow
Directory of Open Access Journals (Sweden)
Dimitra C. Douvi
2017-02-01
Full Text Available This paper opens up a new perspective on the aerodynamic performance of a wind turbine airfoil. More specifically, the paper deals with a steady, incompressible two-phase flow, consisting of air and two different concentrations of sand particles, over an airfoil from the National Renewable Energy Laboratory, NREL S809. The numerical simulations were performed on turbulence models for aerodynamic operations using commercial computational fluid dynamics (CFD code. The computational results obtained for the aerodynamic performance of an S809 airfoil at various angles of attack operating at Reynolds numbers of Re = 1 × 106 and Re = 2 × 106 in a dry, dusty environment were compared with existing experimental data on air flow over an S809 airfoil from reliable sources. Notably, a structured mesh consisting of 80,000 cells had already been identified as the most appropriate for numerical simulations. Finally, it was concluded that sand concentration significantly affected the aerodynamic performance of the airfoil; there was an increase in the values of the predicted drag coefficients, as well as a decrease in the values of the predicted lift coefficients caused by increasing concentrations of sand particles. The region around the airfoil was studied by using contours of static pressure and discrete phase model (DPM concentration.
Jumahadi, Muhammad Taufiq; Saad, Mohd Rashdan; Idris, Azam Che; Sujipto, Suriyadi; Rahman, Mohd Rosdzimin Abdul
2018-02-01
Boundary layer separation is detrimental to the lift and drag of most aeronautical applications. Many vortex generators (VG), both passive and active have been designed to reduce these drawbacks. This study targets to investigate the effectiveness of hybrid micro-VGs, which combine both active and passive micro-VGs in controlling separation under subsonic conditions. NACA 4415 airfoils installed with passive, active and hybrid micro-VGs each are designed, 3D printed, and tested in a wind tunnel at 26.19 m/s under Re = 2.5x105. The lift and drag measurements from a 3-component force balance prove that hybrid micro-VGs increase lift by up to 21.2%, increase drag by more than 11.3% and improve lift-to-drag ratio by at least 8.6% until up to 33.7%. From this research, it is believed that hybrid micro-VGs are competitive to the performance of active VGs and a better configuration is to be considered to reduce parasitic drag and outstand active VGs.
DEFF Research Database (Denmark)
Fischer, Andreas
2011-01-01
The present work aims at the characterization of aerodynamic noise from wind turbines. There is a consensus among scientists that the dominant aerodynamic noise mechanism is turbulent boundary trailing edge noise. In almost all operational conditions the boundary layer flow over the wind turbine...... blades makes a transition from laminar to turbulent. In the turbulent boundary layer eddies are created which are a potential noise sources. They are ineffective as noise source on the airfoil surface or in free flow, but when convecting past the trailing edge of the airfoil their efficiency is much...... for aerodynamic wind tunnels with a hard wall test section. Acoustic far field sound measurements are not possible in this tunnel due to the high background noise. The second wind tunnel is owned by Virginia Tech University. The test section has Kevlar walls which are acoustically transparent and it is surrounded...
URANS simulations of separated flow with stall cells over an NREL S826 airfoil
Sarlak, H.; Nishino, T.; Sørensen, J. N.
2016-06-01
A series of wind tunnel measurements and oil flow visualization was recently carried out at the Technical University of Denmark in order to investigate flow characteristics over a 14% thick NREL S826 airfoil at low Reynolds numbers. This paper aims at presenting numerical simulations of the same airfoil using unsteady Reynolds-averaged Navier-Stokes (URANS) approach. Results of the simulations are demonstrated in terms of mean flow velocity, lift and drag, as well as pressure distribution, and validated against available experimental data. The simulations are carried out with a wide computational domain (with a span-to-chord ratio of 5) and it is illustrated that the URANS approach is capable of predicting 3D spanwise structures, known as stall cells.
CFD aerodynamic analysis of non-conventional airfoil sections for very large rotor blades
Papadakis, G.; Voutsinas, S.; Sieros, G.; Chaviaropoulos, T.
2014-12-01
The aerodynamic performance of flat-back and elliptically shaped airfoils is analyzed on the basis of CFD simulations. Incompressible and low-Mach preconditioned compressible unsteady simulations have been carried out using the k-w SST and the Spalart Allmaras turbulence models. Time averaged lift and drag coefficients are compared to wind tunnel data for the FB 3500-1750 flat back airfoil while amplitudes and frequencies are also recorded. Prior to separation averaged lift is well predicted while drag is overestimated keeping however the trend in the tests. The CFD models considered, predict separation with a 5° delay which is reflected on the load results. Similar results are provided for a modified NACA0035 with a rounded (elliptically shaped) trailing edge. Finally as regards the dynamic characteristics in the load signals, there is fair agreement in terms of Str number but significant differences in terms of lift and drag amplitudes.
Experimental benchmark and code validation for airfoils equipped with passive vortex generators
DEFF Research Database (Denmark)
Baldacchino, D.; Manolesos, M.; Ferreira, Célia Maria Dias
2016-01-01
% thick DU97W300 and an 18% thick NTUA T18 have been used for benchmarking several simulation tools. These tools span low-to-high complexity, ranging from engineering-level integral boundary layer tools to fully-resolved computational fluid dynamics codes. Results indicate that with appropriate......Experimental results and complimentary computations for airfoils with vortex generators are compared in this paper, as part of an effort within the AVATAR project to develop tools for wind turbine blade control devices. Measurements from two airfoils equipped with passive vortex generators, a 30...... calibration, engineering-type tools can capture the effects of vortex generators and outperform more complex tools. Fully resolved CFD comes at a much higher computational cost and does not necessarily capture the increased lift due to the VGs. However, in lieu of the limited experimental data available...
Aeroelastic Stability of a 2D Airfoil Section equipped with a Trailing Edge Flap
DEFF Research Database (Denmark)
Bergami, Leonardo
Recent studies conclude that important reduction of the fatigue loads encountered by a wind turbine blade can be achieved using a deformable trailing edge control system. The focus of the current work is to determine the effect of this flap-like system on the aeroelastic stability of a 2D airfoil...... section. A simulation tool is implemented to predict the flow speed at which a flap equipped section may become unstable, either due to flutter or divergence. First, the stability limits of the airfoil without flap are determined, and, in the second part of the work, a deformable trailing edge flap...... is applied. Stability is investigated for the uncontrolled flap, and for three different control algorithms. The three controls are tuned for fatigue load alleviation and they are based on, respectively, measurement of the heave displacement and velocity, measurement of the local angle of attack, measurement...
Directory of Open Access Journals (Sweden)
Gunawan Nugroho
2012-09-01
Full Text Available Abstract: The Characteristic of Stream Aerodynamics through Airfoil Joukowski/ Flat Plate on Endwall Area. This research analyzed the characteristic of stream aerodynamics through airfoil joukowski/flat plate. It focuses on endwall area due to its complexity of three-dimensional flow and the numerical method as its characteristic. The numerical method is used to solve the experiment peripherals obstacles that have to cover some detailed aspects. The result shows that in zero-degree angle of attack, the saddle point has been formed near the leading edge. The streamline profile seems symmetric in endwall area and shows near-nil lift values. In 10-degree angle of attack, the saddle point has been formed downward and away from leading edge. It reveals that the adverse pressure gradient raises as angle of attack increased, and forces wider low-pressure area to strengthen flow from span to endwall area.
Large-eddy simulation of flow around an airfoil on a structured mesh
Kaltenbach, Hans-Jakob; Choi, Haecheon
1995-01-01
The diversity of flow characteristics encountered in a flow over an airfoil near maximum lift taxes the presently available statistical turbulence models. This work describes our first attempt to apply the technique of large-eddy simulation to a flow of aeronautical interest. The challenge for this simulation comes from the high Reynolds number of the flow as well as the variety of flow regimes encountered, including a thin laminar boundary layer at the nose, transition, boundary layer growth under adverse pressure gradient, incipient separation near the trailing edge, and merging of two shear layers at the trailing edge. The flow configuration chosen is a NACA 4412 airfoil near maximum lift. The corresponding angle of attack was determined independently by Wadcock (1987) and Hastings & Williams (1984, 1987) to be close to 12 deg. The simulation matches the chord Reynolds number U(sub infinity)c/v = 1.64 x 10(exp 6) of Wadcock's experiment.
Modelling of unsteady airfoil aerodynamics for the prediction of blade standstill vibrations
DEFF Research Database (Denmark)
Skrzypinski, Witold Robert; Gaunaa, Mac; Sørensen, Niels N.
2012-01-01
In the present work, CFD simulations of the DU96-W-180 airfoil at 26 and 24 deg. angles of attack were performed. 2D RANS and 3D DES computations with non-moving and prescribed motion airfoil suspensions were carried out. The openings of the lift coefficient loops predicted by CFD were different...... that further investigations are needed and that caution should be taken when applying engineering models in connection with aeroelastic simulations. Nonetheless, the results of the 2D CFD, 3D CFD and the engineering models indicate that the associated aerodynamic damping may be higher than that predicted...... than those predicted by engineering models. The average lift slope of the loops from the 3D CFD had opposite sign than the one from 2D CFD. Trying to model the 3D behaviour with the engineering models proved difficult. The disagreement between the 2D CFD, 3D CFD and the engineering models indicates...
Comparative computational analysis of airfoil sections for use on sailing craft
Partida, Louis P.
1996-01-01
Approved for public release; distribution is unlimited This thesis represents the results of a comparative analysis of current and proposed airfoil sections for use on sailing craft. The primary goal of this report is to develop a sail replacement that functions with the ease and durability of current sailboat sails, yet offers a marked improvement in overall performance, with minimum penalties of weight and construction complexity. State-of-the-art computational methods are utilized to de...
Numerical simulation of aerodynamic sound radiated from a two-dimensional airfoil
飯田, 明由; 大田黒, 俊夫; 加藤, 千幸; Akiyoshi, Iida; Toshio, Otaguro; Chisachi, Kato; 日立機研; 日立機研; 東大生研; Mechanical Engineering Research Laboratory, Hitachi Ltd.; Mechanical Engineering Research Laboratory, Hitachi Ltd.; University of Tokyo
2000-01-01
An aerodynamic sound radiated from a two-dimensional airfoil has been computed with the Lighthill-Curle's theory. The predicted sound pressure level is agreement with the measured one. Distribution of vortex sound sources is also estimated based on the correlation between the unsteady vorticity fluctuations and the aerodynamic sound. The distribution of vortex sound source reveals that separated shear layers generate aerodynamic sound. This result is help to understand noise reduction method....
Computing Aerodynamic Performance of a 2D Iced Airfoil: Blocking Topology and Grid Generation
Chi, X.; Zhu, B.; Shih, T. I.-P.; Slater, J. W.; Addy, H. E.; Choo, Yung K.; Lee, Chi-Ming (Technical Monitor)
2002-01-01
The ice accrued on airfoils can have enormously complicated shapes with multiple protruded horns and feathers. In this paper, several blocking topologies are proposed and evaluated on their ability to produce high-quality structured multi-block grid systems. A transition layer grid is introduced to ensure that jaggedness on the ice-surface geometry do not to propagate into the domain. This is important for grid-generation methods based on hyperbolic PDEs (Partial Differential Equations) and algebraic transfinite interpolation. A 'thick' wrap-around grid is introduced to ensure that grid lines clustered next to solid walls do not propagate as streaks of tightly packed grid lines into the interior of the domain along block boundaries. For ice shapes that are not too complicated, a method is presented for generating high-quality single-block grids. To demonstrate the usefulness of the methods developed, grids and CFD solutions were generated for two iced airfoils: the NLF0414 airfoil with and without the 623-ice shape and the B575/767 airfoil with and without the 145m-ice shape. To validate the computations, the computed lift coefficients as a function of angle of attack were compared with available experimental data. The ice shapes and the blocking topologies were prepared by NASA Glenn's SmaggIce software. The grid systems were generated by using a four-boundary method based on Hermite interpolation with controls on clustering, orthogonality next to walls, and C continuity across block boundaries. The flow was modeled by the ensemble-averaged compressible Navier-Stokes equations, closed by the shear-stress transport turbulence model in which the integration is to the wall. All solutions were generated by using the NPARC WIND code.
Investigation of unsteady flow development over a pitching airfoil by means of TR-PIV
Mulleners, Karen; Henning, Arne; Mai, Holger; Raffel, Markus; Costes, Michel; Le Pape, Arnaud
2009-01-01
The flow over an OA209 airfoil subjected to a sinusoidal pitching motion under dynamic stall conditions is investigated experimentally by means of time resolved particle image velocimetry (TR-PIV) and surface pressure measurements. Dynamic stall is distinguished by the formation and convection of large scale coherent structures and a delay in massive flow separation. A vortex detection scheme based on an identification function derived directly from the velocity fields is adopted to identify ...
Vets, Robert
An experimental study was conducted to assess the application of a moving surface to affect boundary layers and circulation around airfoils for the purpose of altering and enhancing aerodynamic performance of finite wings at moderate Reynolds numbers. The moving surface was established by a wide, lightweight, nylon belt that enveloped a wing's symmetric airfoil profile articulated via a friction drive cylinder such that the direction of the upper surface was in the direction of the free stream. A water tunnel visualization study accompanied wind tunnel testing at the University of Washington, Kirsten Wind Tunnel of finite wings. An experimental study was conducted to assess the application of a moving surface to affect boundary layers and circulation around airfoils for the purpose of altering and enhancing aerodynamic performance of finite wings at moderate Reynolds numbers. The moving surface was established by a wide, lightweight, nylon belt that enveloped a wing's symmetric airfoil profile articulated via a friction drive cylinder such that the direction of the upper surface was in the direction of the free stream. A water tunnel visualization study accompanied wind tunnel testing at the University of Washington, Kirsten Wind Tunnel of finite wings. The defining non-dimensional parameter for the system is the ratio of the surface velocity to the free stream velocity, us/Uo. Results show a general increase in lift with increasing us/Uo. The endurance parameter served as an additional metric for the system's performance. Examining the results of the endurance parameter shows general increase in endurance and lift with the moving surface activated. Peak performance in terms of increased endurance along with increased lift occurs at or slightly above us/Uo = 1. Water tunnel visualization showed a marked difference in the downwash for velocity ratios greater than 1, supporting the measured data. Reynolds numbers for this investigation were 1.9E5 and 4.3E5, relevant
Numerical solution of inviscid and viscous laminar and turbulent flow around the airfoil
Directory of Open Access Journals (Sweden)
Slouka Martin
2016-01-01
Full Text Available This work deals with the 2D numerical solution of inviscid compressible flow and viscous compressible laminar and turbulent flow around the profile. In a case of turbulent flow algebraic Baldwin-Lomax model is used and compared with Wilcox k-omega model. Calculations are done for NACA 0012 and RAE 2822 airfoil profile for the different angles of upstream flow. Numerical results are compared and discussed with experimental data.
Numerical solution of inviscid and viscous laminar and turbulent flow around the airfoil
Slouka, Martin; Kozel, Karel
2016-03-01
This work deals with the 2D numerical solution of inviscid compressible flow and viscous compressible laminar and turbulent flow around the profile. In a case of turbulent flow algebraic Baldwin-Lomax model is used and compared with Wilcox k-omega model. Calculations are done for NACA 0012 and RAE 2822 airfoil profile for the different angles of upstream flow. Numerical results are compared and discussed with experimental data.
On the effects of leading edge vortex generators on an OA209 airfoil
Heine, Benjamin; Mulleners, Karen; Gardner, Anthony; Mai, Holger
2009-01-01
Leading edge vortex generators have been found to significantly increase the aerodynamic performance of an airfoil under dynamic stall conditions. However, the principle of operation of these devices is still unclear. Therefore static wind and water tunnel experiments as well as CFD simulations have been conducted on a rotary aircraft wing profile OA209. A POD analysis applied to the vector fields generated by PIV measurements showed that the vortex generators break larger flow structures...
Knowledge-based system for detailed blade design of turbines
Goel, Sanjay; Lamson, Scott
1994-03-01
A design optimization methodology that couples optimization techniques to CFD analysis for design of airfoils is presented. This technique optimizes 2D airfoil sections of a blade by minimizing the deviation of the actual Mach number distribution on the blade surface from a smooth fit of the distribution. The airfoil is not reverse engineered by specification of a precise distribution of the desired Mach number plot, only general desired characteristics of the distribution are specified for the design. Since the Mach number distribution is very complex, and cannot be conveniently represented by a single polynomial, it is partitioned into segments, each of which is characterized by a different order polynomial. The sum of the deviation of all the segments is minimized during optimization. To make intelligent changes to the airfoil geometry, it needs to be associated with features observed in the Mach number distribution. Associating the geometry parameters with independent features of the distribution is a fairly complex task. Also, for different optimization techniques to work efficiently the airfoil geometry needs to be parameterized into independent parameters, with enough degrees of freedom for adequate geometry manipulation. A high-pressure, low reaction steam turbine blade section was optimized using this methodology. The Mach number distribution was partitioned into pressure and suction surfaces and the suction surface distribution was further subdivided into leading edge, mid section and trailing edge sections. Two different airfoil representation schemes were used for defining the design variables of the optimization problem. The optimization was performed by using a combination of heuristic search and numerical optimization. The optimization results for the two schemes are discussed in the paper. The results are also compared to a manual design improvement study conducted independently by an experienced airfoil designer. The turbine blade optimization
Evaluation of tunnel sidewall boundary-layer-control systems for high-lift airfoil testing
Paschal, K.; Goodman, W.; Mcghee, R.; Walker, B.; Wilcox, Peter A.
1991-01-01
An experimental study was conducted in the NASA Langley Low-Turbulence Pressure Tunnel to evaluate a suction sidewall boundary-layer-control (BLC) technique used in testing 2D high-lift airfoils. Sidewall BLC is required to maintain spanwise two-dimensionality of the flow over the airfoil at large angles of attack. A supercritical-type high-lift air-foil, equipped with a double-slotted flap and a leading-edge slat, was used for the study which was conducted at a Mach number of 0.2 and Reynolds numbers based on chord of 9 and 16 million. The sidewall BLC technique, which features distributed suction through porous endplates connected to a venting system, was able to control sidewall boundary-layer separation and maintain two-dimensional flow over the high-lift configuration for both Reynolds numbers tested. Discussions on porous endplate optimization and effects of suction on section lift are presented. Results obtained with the suction system were also compared with previous data obtained with a tangential blowing BLC system for the same high-lift configuration.
Modeling of tangential synthetic jet actuators used for pitching control on an airfoil
Lopez, Omar; Moser, Robert
2008-11-01
Pitching moment control in an airfoil can be achieved by trapping concentrations of vorticity close to the trailing edge. Experimental work has shown that synthetic jet actuators can be used to manipulate and control this trapped vorticity. Two different approaches are used to model the action of tangential-blowing synthetic jet actuators mounted near the trailing edge of the airfoil: a detailed model and Reynolds stress synthetic jet (RSSJ) model. The detailed model resolves the synthetic jet dynamics in time while the RSSJ model tries to capture the major effects of the synthetic jet by modeling the changes in the Reynolds stress induced by the actuator, based on experimental PIV data and numerical results from the detailed model. Both models along with the CFD computations in which they are embedded are validated against experimental data. The synthetic jet models have been developed to simulate closed loop flow control of the pitching and plunging of the airfoil, and to this end the RSSJ model is particularly useful since it reduces (by an order of magnitude) the cost of simulating the long-term evolution of the system under control.
Boundary Layer Separation and Reattachment Detection on Airfoils by Thermal Flow Sensors
Directory of Open Access Journals (Sweden)
Peter Busche
2012-10-01
Full Text Available A sensor concept for detection of boundary layer separation (flow separation, stall and reattachment on airfoils is introduced in this paper. Boundary layer separation and reattachment are phenomena of fluid mechanics showing characteristics of extinction and even inversion of the flow velocity on an overflowed surface. The flow sensor used in this work is able to measure the flow velocity in terms of direction and quantity at the sensor’s position and expected to determine those specific flow conditions. Therefore, an array of thermal flow sensors has been integrated (flush-mounted on an airfoil and placed in a wind tunnel for measurement. Sensor signals have been recorded at different wind speeds and angles of attack for different positions on the airfoil. The sensors used here are based on the change of temperature distribution on a membrane (calorimetric principle. Thermopiles are used as temperature sensors in this approach offering a baseline free sensor signal, which is favorable for measurements at zero flow. Measurement results show clear separation points (zero flow and even negative flow values (back flow for all sensor positions. In addition to standard silicon-based flow sensors, a polymer-based flexible approach has been tested showing similar results.
Zonal Detached-Eddy Simulation of Turbulent Unsteady Flow over Iced Airfoils
Zhang, Yue
2015-07-23
This paper presentsamultiscale finite-element formulation for the second modeofzonal detached-eddy simulation. The multiscale formulation corrects the lack of stability of the standard Galerkin formulation by incorporating the effect of unresolved scales to the grid (resolved) scales. The stabilization terms arise naturally and are free of userdefined stability parameters. Validation of the method is accomplished via the turbulent flow over tandem cylinders. The boundary-layer separation, free shear-layer rollup, vortex shedding from the upstream cylinder, and interaction with the downstream cylinder are well reproduced. Good agreement with experimental measurements gives credence to the accuracy of zonal detached-eddy simulation in modeling turbulent separated flows. A comprehensive study is then conducted on the performance degradation of ice-contaminated airfoils. NACA 23012 airfoil with a spanwise ice ridge and Gates Learjet Corporation-305 airfoil with a leading-edge horn-shape glaze ice are selected for investigation. Appropriate spanwise domain size and sufficient grid density are determined to enhance the reliability of the simulations. A comparison of lift coefficient and flowfield variables demonstrates the added advantage that the zonal detached-eddy simulation model brings to the Spalart-Allmaras turbulence model. Spectral analysis and instantaneous visualization of turbulent structures are also highlighted via zonal detached-eddy simulation. Copyright © 2015 by the CFD Lab of McGill University. Published by the American Institute of Aeronautics and Astronautics, Inc.
LPV Modeling and Control for Active Flutter Suppression of a Smart Airfoil
Al-Hajjar, Ali M. H.; Al-Jiboory, Ali Khudhair; Swei, Sean Shan-Min; Zhu, Guoming
2018-01-01
In this paper, a novel technique of linear parameter varying (LPV) modeling and control of a smart airfoil for active flutter suppression is proposed, where the smart airfoil has a groove along its chord and contains a moving mass that is used to control the airfoil pitching and plunging motions. The new LPV modeling technique is proposed that uses mass position as a scheduling parameter to describe the physical constraint of the moving mass, in addition the hard constraint at the boundaries is realized by proper selection of the parameter varying function. Therefore, the position of the moving mass and the free stream airspeed are considered the scheduling parameters in the study. A state-feedback based LPV gain-scheduling controller with guaranteed H infinity performance is presented by utilizing the dynamics of the moving mass as scheduling parameter at a given airspeed. The numerical simulations demonstrate the effectiveness of the proposed LPV control architecture by significantly improving the performance while reducing the control effort.
Controlled aeroelastic response and airfoil shaping using adaptive materials and integrated systems
Pinkerton, Jennifer L.; McGowan, Anna-Maria R.; Moses, Robert W.; Scott, Robert C.; Heeg, Jennifer
1996-05-01
This paper presents an overview of several activities of the Aeroelasticity Branch at the NASA Langley Research Center in the area of applying adaptive materials and integrated systems for controlling both aircraft aeroelastic response and airfoil shape. The experimental results of four programs are discussed: the Piezoelectric Aeroelastic Response Tailoring Investigation (PARTI); the adaptive neural control of aeroelastic response (ANCAR) program; the actively controlled response of buffet affected tails (ACROBAT) program; and the Airfoil THUNDER Testing to ascertain charcteristics (ATTACH) project. The PARTI program demonstrated active flutter control and significant reductions in aeroelastic response at dynamic pressures below flutter using piezoelectric actuators. The ANCAR program seeks to demonstrate the effectiveness of using neural networks to schedule flutter suppression control laws. The ACROBAT program studied the effectiveness of a number of candidate actuators, including a rudder and piezoelectric actuators, to alleviate vertical tail buffeting. In the ATTACH project, the feasibility of using thin-layer composite-unimorph piezoelectric driver and sensor (THUNDER) wafers to control airfoil aerodynamic characteristics was investigated. Plans for future applications are also discussed.
Reynolds number effect on airfoil wake structures under pitching and heaving motion
Kim, Kyung Chun; Karbasian, Hamidreza; ExpTENsys Team
2017-11-01
Detached Eddy Simulation (DES) and particle image velocimetry (PIV) measurements were performed to investigate the wake flow characteristics of an airfoil under pitching and heaving motion. A NACA0012 airfoil was selected for the numerical simulation and experiments were carried out in a wind tunnel and a water tunnel at Reynolds number of 15,000 and 90,000, respectively. The airfoil oscillated around an axis located 1/4 distance from the leading edge chord. Two different angles of attack, 20° and 30°, were selected with +/-10° maximum amplitude of oscillation. In order to extract the coherent flow structures from time-resolved PIV data, proper orthogonal decomposition (POD) analysis was performed on 1,000 instantaneous realisations for each condition using the method of snapshots. Vorticity contour and velocity profiles for both PIV and DES results are in good agreement for pitching and heaving motion. At high Reynolds number, 3D stream-wise vortices appeared after generating span-wise vortices. The higher maximum angle of attack allows the leading edge vortex to grow stronger and that the angle of attack appears to be more important in influencing the growth of the leading edge vortex structure than the reduced frequency. National Research Foundation of Korea (No. 2011-0030013).
Effects of leading and trailing edge flaps on the aerodynamics of airfoil/vortex interactions
Hassan, Ahmed A.; Sankar, L. N.; Tadghighi, H.
1994-01-01
A numerical procedure has been developed for predicting the two-dimensional parallel interaction between a free convecting vortex and a NACA 0012 airfoil having leading and trailing edge integral-type flaps. Special emphasis is placed on the unsteady flap motion effects which result in alleviating the interaction at subcritical and supercritical onset flows. The numerical procedure described here is based on the implicit finite-difference solutions to the unsteady two-dimensional full potential equation. Vortex-induced effects are computed using the Biot-Savart Law with allowance for a finite core radius. The vortex-induced velocities at the surface of the airfoil are incorporated into the potential flow model via the use of the velocity transpiration approach. Flap motion effects are also modeled using the transpiration approach. For subcritical interactions, our results indicate that trailing edge flaps can be used to alleviate the impulsive loads experienced by the airfoil. For supercritical interactions, our results demonstrate the necessity of using a leading edge flap, rather than a trailing edge flap, to alleviate the interaction. Results for various time-dependent flap motions and their effect on the predicted temporal sectional loads, differential pressures, and the free vortex trajectories are presented
A CFD Database for Airfoils and Wings at Post-Stall Angles of Attack
Petrilli, Justin; Paul, Ryan; Gopalarathnam, Ashok; Frink, Neal T.
2013-01-01
This paper presents selected results from an ongoing effort to develop an aerodynamic database from Reynolds-Averaged Navier-Stokes (RANS) computational analysis of airfoils and wings at stall and post-stall angles of attack. The data obtained from this effort will be used for validation and refinement of a low-order post-stall prediction method developed at NCSU, and to fill existing gaps in high angle of attack data in the literature. Such data could have potential applications in post-stall flight dynamics, helicopter aerodynamics and wind turbine aerodynamics. An overview of the NASA TetrUSS CFD package used for the RANS computational approach is presented. Detailed results for three airfoils are presented to compare their stall and post-stall behavior. The results for finite wings at stall and post-stall conditions focus on the effects of taper-ratio and sweep angle, with particular attention to whether the sectional flows can be approximated using two-dimensional flow over a stalled airfoil. While this approximation seems reasonable for unswept wings even at post-stall conditions, significant spanwise flow on stalled swept wings preclude the use of two-dimensional data to model sectional flows on swept wings. Thus, further effort is needed in low-order aerodynamic modeling of swept wings at stalled conditions.
Guda, Venkata Subba Sai Satish
There have been several advancements in the aerospace industry in areas of design such as aerodynamics, designs, controls and propulsion; all aimed at one common goal i.e. increasing efficiency --range and scope of operation with lesser fuel consumption. Several methods of flow control have been tried. Some were successful, some failed and many were termed as impractical. The low Reynolds number regime of 104 - 105 is a very interesting range. Flow physics in this range are quite different than those of higher Reynolds number range. Mid and high altitude UAV's, MAV's, sailplanes, jet engine fan blades, inboard helicopter rotor blades and wind turbine rotors are some of the aerodynamic applications that fall in this range. The current study deals with using dynamic roughness as a means of flow control over a NACA 0012 airfoil at low Reynolds numbers. Dynamic 3-D surface roughness elements on an airfoil placed near the leading edge aim at increasing the efficiency by suppressing the effects of leading edge separation like leading edge stall by delaying or totally eliminating flow separation. A numerical study of the above method has been carried out by means of a Large Eddy Simulation, a mathematical model for turbulence in Computational Fluid Dynamics, owing to the highly unsteady nature of the flow. A user defined function has been developed for the 3-D dynamic roughness element motion. Results from simulations have been compared to those from experimental PIV data. Large eddy simulations have relatively well captured the leading edge stall. For the clean cases, i.e. with the DR not actuated, the LES was able to reproduce experimental results in a reasonable fashion. However DR simulation results show that it fails to reattach the flow and suppress flow separation compared to experiments. Several novel techniques of grid design and hump creation are introduced through this study.
Application of FLEET Velocimetry in the NASA Langley 0.3-meter Transonic Cryogenic Tunnel
Burns, Ross A.; Danehy, Paul M.; Halls, Benjamin R.; Jiang, Naibo
2015-01-01
Femtosecond laser electronic excitation and tagging (FLEET) velocimetry is demonstrated in a large-scale transonic cryogenic wind tunnel. Test conditions include total pressures, total temperatures, and Mach numbers ranging from 15 to 58 psia, 200 to 295 K, and 0.2 to 0.75, respectively. Freestream velocity measurements exhibit accuracies within 1 percent and precisions better than 1 m/s. The measured velocities adhere closely to isentropic flow theory over the domain of temperatures and pressures that were tested. Additional velocity measurements are made within the tunnel boundary layer; virtual trajectories traced out by the FLEET signal are indicative of the characteristic turbulent behavior in this region of the flow, where the unsteadiness increases demonstrably as the wall is approached. Mean velocities taken within the boundary layer are in agreement with theoretical velocity profiles, though the fluctuating velocities exhibit a greater deviation from theoretical predictions.
Transonic wind tunnel tests of a .015 scale space shuttle orbiter model, volume 2
Struzynski, N. A.
1975-01-01
Transonic wind tunnel tests were run on a 0.015 scale model of the Space Shuttle Orbiter Vehicle in an eight-foot tunnel during August 1975. The purpose of the program was to obtain basic shuttle aerodynamic data through a full range of elevon and aileron deflections, verification of data obtained at other facilities, and effects of Reynolds numbers. The second part of a discussion of test procedures and results in both tabular and graphical form were presented. Tests were performed at Mach numbers from 0.35 to 1.20, and at Reynolds numbers from 3.5 million to 8.2 million per foot. The angle of attack was varied from -2 to +20 degrees at sideslip angles of -2, 0, +2 degrees. Sideslip was varied from -6 to +8 degrees at constant angles of attack from 0 to +20 degrees. Various aileron and ailevon settings were tested for various angles of attack.
Transonic wind tunnel tests of A.015 scale space shuttle orbiter model, volume 1
Struzynski, N. A.
1975-01-01
Transonic wind tunnel tests were run on a 0.015 scale model of the Space Shuttle Orbiter Vehicle in an eight-foot tunnel during August 1975. The purpose of the program was to obtain basic shuttle aerodynamic data through a full range of elevon and aileron deflections, verification of data obtained at other facilities, and effects of Reynolds numbers. The first part of a discussion of test procedures and results in both tabular and graphical form were presented. Tests were performed at Mach numbers from 0.35 to 1.20, and at Reynolds numbers for 3.5 million to 8.2 million per foot. The angle of attack was varied from -1 to +20 degrees at sideslip angles of -2, 0, +2 degrees. Sideslip was varied from -6 to +8 degrees at constant angles of attack from 0 to +20 degrees. Various aileron and ailevon settings were tested for various angles of attack.
A finite element formulation of Euler equations for the solution of steady transonic flows
Ecer, A.; Akay, H. U.
1982-01-01
The main objective of the considered investigation is related to the development of a relaxation scheme for the analysis of inviscid, rotational, transonic flow problems. To formulate the equations of motion for inviscid flows in a fixed coordinate system, an Eulerian type variational principle is required. The derivation of an Eulerian variational principle which is employed in the finite element formulation is discussed. The presented numerical method describes the mathematical formulation and the application of a numerical process for the direct solution of steady Euler equations. The development of the procedure as an extension of existing potential flow formulations provides the applicability of previous procedures, e.g., proper application of the artificial viscosity for supersonic elements, and the accurate modeling of the shock.
Control of the NASA Langley 16-Foot Transonic Tunnel with the Self-Organizing Map
Motter, Mark A.
1999-01-01
A predictive, multiple model control strategy is developed based on an ensemble of local linear models of the nonlinear system dynamics for a transonic wind tunnel. The local linear models are estimated directly from the weights of a self-organizing map (SOM). Multiple self-organizing maps collectively model the global response of the wind tunnel to a finite set of representative prototype controls. These prototype controls partition the control space and incorporate experiential knowledge gained from decades of operation. Each SOM models the combination of the tunnel with one of the representative controls, over the entire range of operation. The SOM based linear models are used to predict the tunnel response to a larger family of control sequences which are clustered on the representative prototypes. The control sequence which corresponds to the prediction that best satisfies the requirements on the system output is applied as the external driving signal.
Computational Analysis of the Transonic Dynamics Tunnel Using FUN3D
Energy Technology Data Exchange (ETDEWEB)
Chwalowski, Pawel; Quon, Eliot; Brynildsen, Scott E.
2016-01-04
This paper presents results from an explanatory two-year effort of applying Computational Fluid Dynamics (CFD) to analyze the empty-tunnel flow in the NASA Langley Research Center Transonic Dynamics Tunnel (TDT). The TDT is a continuous-flow, closed circuit, 16- x 16-foot slotted-test-section wind tunnel, with capabilities to use air or heavy gas as a working fluid. In this study, experimental data acquired in the empty tunnel using the R-134a test medium was used to calibrate the computational data. The experimental calibration data includes wall pressures, boundary-layer profiles, and the tunnel centerline Mach number profiles. Subsonic and supersonic flow regimes were considered, focusing on Mach 0.5, 0.7 and Mach 1.1 in the TDT test section. This study discusses the computational domain, boundary conditions, and initial conditions selected in the resulting steady-state analyses using NASA's FUN3D CFD software.
Hsieh, T.; Coakley, T. J.
1987-01-01
An investigation of downstream boundary effects on the frequency of self-excited oscillations in two-dimensional, separated transonic diffuser flows has been conducted numerically by solving the compressible, Reynolds-averaged, thin-layer Navier-Stokes equation with a two-equation turbulence model. It was found that the unsteady diffuser flowfields are very sensitive to the location of the downstream boundary. Extension of the diffuser downstream boundary significantly reduces the frequency and amplitude of oscillations for pressure, velocity and shock. Computational results suggest that the mechanism causing the self-excited oscillation changes from viscous convective wave dominated oscillations to inviscid acoustic wave dominated oscillations when the location of downstream boundary varies from 8.66 to 134.7 throat height. The existence of a suction slot in the experimental setup obscures the physical downstream boundary and, therefore, presents a difficulty for quantitative comparisons between computation and experiment.
Hurley, G. F.; Marr, N. W.
1975-01-01
An attempt was made to produce airfoil shaped bars of three different eutectic superalloys by means of the edge-defined, film-fed growth (EFG) method. The alloys used were a gamma + delta Ni-Cb alloy, a gamma/gamma prime + delta Ni-Cb-Al alloy and a Co-TaC alloy containing Ni and Cr. The development of a new die material was essential in the investigation since these alloys are reactive toward known die materials. Tantalum carbide was selected as a die material because it exhibited spontaneous capillary rise and slow rate of degradation in the liquid metals. Eutectic bars up to 1 mm thick and 6 mm wide were grown from TaC dies in order to determine the growth characteristics and the thermal gradient. Large bars of the gamma/gamma prime + delta alloy were grown and tensile tested. A die with a blind central cavity was designed and several hollow, tear-shaped bars were grown.
Tan, Yuehan; Crittenden, Thomas; Glezer, Ari
2017-11-01
The aerodynamic loads on an airfoil moving in coupled, time-periodic pitch-plunge beyond the static stall margin are controlled using transitory regulation of trapped vorticity concentrations. Actuation is effected by a spanwise array of integrated miniature chemical (combustion based) impulse actuators that are triggered intermittently during the airfoil's motion and have a characteristic time scale that is an order of magnitude shorter than the airfoil's convective time scale. Each actuation pulse effects momentary interruption and suspension of the vorticity flux with sufficient control authority to alter the airfoil's global aerodynamic characteristics throughout its motion cycle. The effects of the actuation are assessed using time-dependent measurements of the lift and pitching moment coupled with time-resolved particle image velocimetry over the airfoil and in its near wake that is acquired phased-locked to its motion. It is shown that while the presence of the pitch-coupled plunge delays lift and moment stall during upstroke, it also delays flow reattachment during the downstroke and results in significant degradation of the pitch stability. These aerodynamic shortcomings are mitigated using superposition of a limited number of pulses that are staged during the pitch/plunge cycle and lead to enhancement of cycle lift and pitch stability, and reduces the cycle hysteresis and peak pitching moment.
DEFF Research Database (Denmark)
Volf, Mette
This publication is unique in its demystification and operationalization of the complex and elusive nature of the design process. The publication portrays the designer’s daily work and the creative process, which the designer is a part of. Apart from displaying the designer’s work methods...... and design parameters, the publication shows examples from renowned Danish design firms. Through these examples the reader gets an insight into the designer’s reality....
International Nuclear Information System (INIS)
Dinh, Cong-Truong; Ma, Sang-Bum; Kim, Kwang Yong
2017-01-01
In this study, stator shroud injection in a single-stage transonic axial compressor is proposed. A parametric study of the effect of stator shroud injection on aerodynamic performances was conducted using the three-dimensional Reynolds-averaged Navier-Stokes equations. The curvature, length, width, and circumferential angle of the stator shroud injector and the air injection mass flow rate were selected as the test parameters. The results of the parametric study show that the aerodynamic performances of the single-stage transonic axial compressor were improved by stator shroud injection. The aerodynamic performances were the most sensitive to the injection mass flow rate. Further, the total pressure ratio and adiabatic efficiency were the maximum when the ratio of circumferential angle was 10%.
Silva, Walter A.; Bennett, Robert M.
1992-01-01
The Computational Aeroelasticity Program-Transonic Small Disturbance (CAP-TSD) code, developed at LaRC, is applied to the active flexible wing wind-tunnel model for prediction of transonic aeroelastic behavior. A semi-span computational model is used for evaluation of symmetric motions, and a full-span model is used for evaluation of antisymmetric motions, and a full-span model is used for evaluation of antisymmetric motions. Static aeroelastic solutions using CAP-TSD are computed. Dynamic deformations are presented as flutter boundaries in terms of Mach number and dynamic pressure. Flutter boundaries that take into account modal refinements, vorticity and entropy corrections, antisymmetric motion, and sensitivity to the modeling of the wing tip ballast stores are also presented with experimental flutter results.
Erickson, Gary E.; Rogers, Lawrence W.
1987-01-01
A subsonic and transonic investigation of the vortex flow behavior of a generic fighter configuration with 55-deg cropped delta wing has been conducted in order to improve current understanding of vortical motions on a wing with deflected leading edge flap at moderate and high angles-of-attack. The leading edge vortex strength was reduced, and the vortex was flatter and closer to the wing surface, as the Mach number increased. Transonically, at high angles-of-attack, the test data suggested the development of a cross-flow shock wave above the vortex sheet which coexisted with a rear shock wave. Subsonically, a deflected leading edge flap was able to sustain a concentrated vortex on the forward-facing surface.
Energy Technology Data Exchange (ETDEWEB)
Dinh, Cong-Truong; Ma, Sang-Bum; Kim, Kwang Yong [Inha Univ., Incheon (Korea, Republic of)
2017-01-15
In this study, stator shroud injection in a single-stage transonic axial compressor is proposed. A parametric study of the effect of stator shroud injection on aerodynamic performances was conducted using the three-dimensional Reynolds-averaged Navier-Stokes equations. The curvature, length, width, and circumferential angle of the stator shroud injector and the air injection mass flow rate were selected as the test parameters. The results of the parametric study show that the aerodynamic performances of the single-stage transonic axial compressor were improved by stator shroud injection. The aerodynamic performances were the most sensitive to the injection mass flow rate. Further, the total pressure ratio and adiabatic efficiency were the maximum when the ratio of circumferential angle was 10%.
Erickson, Gary E.
1991-01-01
The vortex dominated aerodynamic characteristics of a generic 65 degree cropped delta wing model were studied in a wind tunnel at subsonic through supersonic speeds. The lee-side flow fields over the wing-alone configuration and the wing with leading edge extension (LEX) added were observed at M (infinity) equals 0.40 to 1.60 using a laser vapor screen technique. These results were correlated with surface streamline patterns, upper surface static pressure distributions, and six-component forces and moments. The wing-alone exhibited vortex breakdown and asymmetry of the breakdown location at the subsonic and transonic speeds. An earlier onset of vortex breakdown over the wing occurred at transonic speeds due to the interaction of the leading edge vortex with the normal shock wave. The development of a shock wave between the vortex and wing surface caused an early separation of the secondary boundary layer. With the LEX installed, wing vortex breakdown asymmetry did not occur up to the maximum angle of attack in the present test of 24 degrees. The favorable interaction of the LEX vortex with the wing flow field reduced the effects of shock waves on the wing primary and secondary vortical flows. The direct interaction of the wing and LEX vortex cores diminished with increasing Mach number. The maximum attainable vortex-induced pressure signatures were constrained by the vacuum pressure limit at the transonic and supersonic speeds.
Low-Order Modeling of Dynamic Stall on Airfoils in Incompressible Flow
Narsipur, Shreyas
Unsteady aerodynamics has been a topic of research since the late 1930's and has increased in popularity among researchers studying dynamic stall in helicopters, insect/bird flight, micro air vehicles, wind-turbine aerodynamics, and ow-energy harvesting devices. Several experimental and computational studies have helped researchers gain a good understanding of the unsteady ow phenomena, but have proved to be expensive and time-intensive for rapid design and analysis purposes. Since the early 1970's, the push to develop low-order models to solve unsteady ow problems has resulted in several semi-empirical models capable of effectively analyzing unsteady aerodynamics in a fraction of the time required by high-order methods. However, due to the various complexities associated with time-dependent flows, several empirical constants and curve fits derived from existing experimental and computational results are required by the semi-empirical models to be an effective analysis tool. The aim of the current work is to develop a low-order model capable of simulating incompressible dynamic-stall type ow problems with a focus on accurately modeling the unsteady ow physics with the aim of reducing empirical dependencies. The lumped-vortex-element (LVE) algorithm is used as the baseline unsteady inviscid model to which augmentations are applied to model unsteady viscous effects. The current research is divided into two phases. The first phase focused on augmentations aimed at modeling pure unsteady trailing-edge boundary-layer separation and stall without leading-edge vortex (LEV) formation. The second phase is targeted at including LEV shedding capabilities to the LVE algorithm and combining with the trailing-edge separation model from phase one to realize a holistic, optimized, and robust low-order dynamic stall model. In phase one, initial augmentations to theory were focused on modeling the effects of steady trailing-edge separation by implementing a non-linear decambering
Sogukpinar, Haci; Bozkurt, Ismail
2018-02-01
Aerodynamic performance of the airfoil plays the most important role to obtain economically maximum efficiency from a wind turbine. Therefore airfoil should have an ideal aerodynamic shape. In this study, aerodynamic simulation of S809 airfoil is conducted and obtained result compared with previously made NASA experimental result and NREL theoretical data. At first, Lift coefficient, lift to drag ratio and pressure coefficient around S809 airfoil are calculated with SST turbulence model, and are compared with experimental and other theoretical data to correlate simulation correctness of the computational approaches. And result indicates good correlation with both experimental and theoretical data. This calculation point out that as the increasing relative velocity, lift to drag ratio increases. Lift to drag ratio attain maximum at the angle around 6 degree and after that starts to decrease again. Comparison shows that CFD code used in this calculation can predict aerodynamic properties of airfoil.
DEFF Research Database (Denmark)
Sørensen, Niels N.
2009-01-01
When predicting the flow over airfoils and rotors, the laminar-turbulent transition process can be important for the aerodynamic performance. Today, the most widespread approach is to use fully turbulent computations, where the transitional process is ignored and the entire boundary layer...... on the wings or airfoils is handled by the turbulence model. The correlation based transition model has lately shown promising results, and the present paper describes the effort of deriving the two non-public empirical correlations of the model to make the model complete. To verify the model it is applied...... to flow over a flat plate, flow over the S809 and the NACA63-415 airfoils, flow over a prolate spheroid at zero and thirty degrees angle of attack, and finally to the NREL Phase VI wind turbine rotor for the zero yaw upwind cases from the NREL/NASA Ames wind tunnel test. Copyright © 2009 John Wiley & Sons...
Directory of Open Access Journals (Sweden)
Edris Yousefi Rad
2017-08-01
Full Text Available In the present research, considering the importance of desirable steam turbine design, improvement of numerical modeling of steam two-phase flows in convergent and divergent channels and the blades of transonic steam turbines has been targeted. The first novelty of this research is the innovative use of combined Convective Upstream Pressure Splitting (CUSP and scalar methods to update the flow properties at each calculation point. In other words, each property (density, temperature, pressure and velocity at each calculation point can be computed from either the CUSP or scalar method, depending on the least deviation criterion. For this reason this innovative method is named “hybrid method”. The next novelty of this research is the use of an inverse method alongside the proposed hybrid method to find the amount of the important parameter z in the CUSP method, which is herein referred to as “CUSP’s convergence parameter”. Using a relatively simple computational grid, firstly, five cases with similar conditions to those of the main cases under study in this research with available experimental data were used to obtain the value of z by the Levenberg-Marquardt inverse method. With this innovation, first, an optimum value of z = 2.667 was obtained using the inverse method and then directly used for the main cases considered in the research. Given that the aim is to investigate the two-dimensional, steady state, inviscid and adiabatic modeling of steam nucleating flows in three different nozzle and turbine blade geometries, flow simulation was performed using a relatively simple mesh and the innovative proposed hybrid method (scalar + CUSP, with the desired value of z = 2.667 . A comparison between the results of the hybrid modeling of the three main cases with experimental data showed a very good agreement, even within shock zones, including the condensation shock region, revealing the efficiency of this numerical modeling method innovation
Aeroservoelastic stability of a 2D airfoil section equipped with a trailing edge flap
Energy Technology Data Exchange (ETDEWEB)
Bergami, Leonardo
2008-11-15
Recent studies conclude that important reduction of the fatigue loads encountered by a wind turbine blade can be achieved using a deformable trailing edge control system. The focus of the current work is to determine the effect of this flap-like system on the aeroelastic stability of a 2D airfoil section. A simulation tool is implemented to predict the flow speed at which a flap equipped section may become unstable, either due to flutter or divergence. First, the stability limits of the airfoil without flap are determined, and, in the second part of the work, a deformable trailing edge flap is applied. Stability is investigated for the uncontrolled flap, and for three different control algorithms. The three controls are tuned for fatigue load alleviation and they are based on, respectively, measurement of the heave displacement and velocity, measurement of the local angle of attack, measurement of the pressure difference between the two sides of the airfoil. The stability of the aeroservoelastic system in a defined equilibrium state, and for a given flow speed, is then determined by solving an eigenvalue problem. Results show that the trailing edge control system modifies significantly the stability limits of the section. In the investigated case, increased flutter limits are reported when the elastic flap is left without control, whereas, by applying any of the control algorithms, the flutter velocity is reduced. Nevertheless, only in the heave control case the flutter limit becomes critically close to normal operation flow speeds. Furthermore, a marked dependence of the stability limits on the control gain is also observed and, by tuning the gain parameters, flutter and divergence can be suppressed for flow speed even above the flutter velocity encountered with uncontrolled flap. (author)
Effects of grit roughness and pitch oscillations on the NACA 4415 airfoil
Energy Technology Data Exchange (ETDEWEB)
Hoffmann, M.J.; Reuss Ramsay, R.; Gregorek, G.M. [Ohio State Univ., Columbus, OH (United States)
1996-07-01
A NACA 4415 airfoil model was tested in The Ohio State University Aeronautical and Astronautical Research Laboratory 3 x 5 subsonic wind tunnel under steady state and unsteady conditions. The test defined baseline conditions for steady state angles of attack from {minus}10{degree} to +40{degree} and examined unsteady behavior by oscillating the model about its pitch axis for three mean angles, three frequencies, and two amplitudes. For all cases, Reynolds numbers of 0.75, 1, 1.25, and 1.5 million were used. In addition, these were repeated after the application of leading edge grit roughness (LEGR) to determine contamination effects on the airfoil performance. Steady state results of the NACA 4415 testing at Reynolds number of 1.25 million showed a baseline maximum lift coefficient of 1.30 at 12.3{degree} angle of attack. The application of LEGR reduced the maximum lift coefficient by 20% and increased the 0.0090 minimum drag coefficient value by 62%. The zero lift pitching moment of {minus}0.0967 showed a 13% reduction in magnitude to {minus}0.0842 with LEGR applied. Data were also obtained for two pitch oscillation amplitudes: {+-}5.5{degree} and {+-}10{degree}. The larger amplitude consistently gave a higher maximum lift coefficient than the smaller amplitude, and both unsteady maximum lift coefficients were greater than the steady state values. Stall is delayed on the airfoil while the angle of attack is increasing, thereby causing an increase in maximum lift coefficient. A hysteresis behavior was exhibited for all the unsteady test cases. The hysteresis loops were larger for the higher reduced frequencies and for the larger amplitude oscillations. As in the steady case, the effect of LEGR in the unsteady case was to reduce the lift coefficient at high angles of attack. In addition, with LEGR, the hysteresis behavior persisted into lower angles of attack than for the clean case.
Effects of grit roughness and pitch oscillations on the S810 airfoil
Energy Technology Data Exchange (ETDEWEB)
Ramsay, R.R.; Hoffman, M.J.; Gregorek, G.M. [Ohio State Univ., Columbus, OH (United States)
1996-01-01
An S810 airfoil model was tested in The Ohio State University Aeronautical and Astronautical Research Laboratory 3 x 5 subsonic wind tunnel under steady state and unsteady conditions. The test defined baseline conditions for steady state angles of attack from -20{degrees} to +40{degrees} and examined unsteady behavior by oscillating the model about its pitch axis for three mean angles, three frequencies, and two amplitudes. For all cases, Reynolds numbers of 0.75, 1, 1.25, and 1.5 million were used. In addition, the above conditions were repeated after the application of leading edge grit roughness (LEGR) to determine contamination effects on the airfoil performance. Baseline steady state results of the S810 testing showed a maximum lift coefficient of 1.15 at 15.2{degrees}angle of attack. The application of LEGR reduced the maximum lift coefficient by 12% and increased the 0.0085 minimum drag coefficient value by 88%. The zero lift pitching moment of -0.0286 showed a 16% reduction in magnitude to -0.0241 with LEGR applied. Data were also obtained for two pitch oscillation amplitudes: {plus_minus}5.5{degrees} and {plus_minus}10{degrees}. The larger amplitude consistently gave a higher maximum lift coefficient than the smaller amplitude and both sets of unsteady maximum lift coefficients were greater than the steady state values. Stall was delayed on the airfoil while the angle of attack was increasing, thereby causing an increase in maximum lift coefficient. A hysteresis behavior was exhibited for all the unsteady test cases. The hysteresis loops were larger for the higher reduced frequencies and for the larger amplitude oscillations. In addition to the hysteresis behavior, an unusual feature of these data were a sudden increase in the lift coefficient where the onset of stall was expected. As in the steady case, the effect of LEGR in the unsteady case was to reduce the lift coefficient at high angles of attack.
Effects of grit roughness and pitch oscillations on the S801 airfoil
Energy Technology Data Exchange (ETDEWEB)
Ramsay, R.R.; Hoffman, M.J.; Gregorek, G.M. [Ohio State Univ., Columbus, OH (United States)
1996-01-01
Horizontal axis wind turbine rotors experience unsteady aerodynamics due to wind shear when the rotor is yawed, when rotor blades pass through the support tower wake, and when the wind is gusting. An understanding of this unsteady behavior is necessary to assist in the calculation of rotor performance and loads. The rotors also experience performance degradation due to surface roughness. These surface irregularities are due to the accumulation of insect debris, ice, and the aging process. Wind tunnel studies that examine both the steady and unsteady behavior of airfoils can help define pertinent flow phenomena, and the resultant data can be used to validate analytical computer codes. A S801 airfoil model was tested in The Ohio State University Aeronautical and Astronautical Research Laboratory (OSU/AARL) 3x5 subsonic wind tunnel (3x5) under steady flow and stationary model conditions, as well as with the model undergoing pitch oscillations. To study the possible extent of performance loss due to surface roughness, a standard grit pattern (LEGR) was used to simulate leading edge contamination. After baseline cases were completed, the LEGR was applied for both steady state and model pitch oscillation cases. The Reynolds numbers used for steady state conditions were 0.75, 1, 1.25, and 1.5 million, while the angle of attack ranged from -20{degrees} to +40{degrees}. With the model undergoing pitch oscillations, data were acquired at Reynolds numbers of 0.75, 1, 1.25, and 1.4 million, at frequencies of 0.6, 1.2, and 1.8 Hz. Two sine wave forcing functions were used, {plus_minus} 5.5 {degrees}and {plus_minus} 10{degrees}, at mean angles of attack of 8{degrees} 14{degrees} and 20{degrees} For purposes herein, any reference to unsteady conditions means that the airfoil model was in pitch oscillation about the quarter chord.
Ghaffari, Farhad
1999-01-01
Transonic Euler computations, based on unstructured grid methodology, are performed for a proposed High Speed Civil Transport (HSCT) configuration, designated as the Reference H configuration within the High Speed Research (HSR) Program. The predicted results are correlated with appropriate experimental wind-tunnel data for the baseline configuration with and without control surface deflections for a range of angle of attack at M(sub infinity) = 0.95. Good correlations between the predictions and measured data have been obtained for the longitudinal aerodynamic characteristics of the baseline configuration. The incremental effects in the longitudinal aerodynamic characteristics due to horizontal rail deflections as well as wing leading-edge and trailing-edge flap deflections have also been predicted reasonably well. Computational results and correlations with data are also presented for the lateral and directional stability characteristics for a range of angle of attack at a constant sideslip angle as well as a range of sideslip angles at a constant angle of attack. In addition, the results are presented to assess the computational method performance and convergence characteristics.
Wolf, Stephen W. D.
1988-01-01
The Wall Adjustment Strategy (WAS) software provides successful on-line control of the 2-D flexible walled test section of the Langley 0.3-m Transonic Cryogenic Tunnel. This software package allows the level of operator intervention to be regulated as necessary for research and production type 2-D testing using and Adaptive Wall Test Section (AWTS). The software is designed to accept modification for future requirements, such as 3-D testing, with a minimum of complexity. The WAS software described is an attempt to provide a user friendly package which could be used to control any flexible walled AWTS. Control system constraints influence the details of data transfer, not the data type. Then this entire software package could be used in different control systems, if suitable interface software is available. A complete overview of the software highlights the data flow paths, the modular architecture of the software and the various operating and analysis modes available. A detailed description of the software modules includes listings of the code. A user's manual is provided to explain task generation, operating environment, user options and what to expect at execution.
Falco UAV Low Reynolds Airfoil Design and Testing at Galileo Avionica
National Research Council Canada - National Science Library
Cistriani, Luca
2007-01-01
... stream turbulence and in-flight icing. Wind tunnel tests have been performed for the two dimensional wing section and for a complete UAV aircraft configuration in order to confirm theoretical previsions...
Falco UAV Low Reynolds Airfoil Design and Testing at Galileo Avionica
National Research Council Canada - National Science Library
Cistriani, Luca
2007-01-01
UAV operations are examined from a performance and logistic flexibility point of view in order to set up requirements to be input for the multiobjective optimization of a two component simple rotation...
Design and Modeling of Turbine Airfoils with Active Flow Control in Realistic Engine Conditions
2008-07-16
for cylinders make using a simple 2d model less meaningful. The solver used for the cylinder cases was SFELES, a quasi 3D large eddy simulation that...would take into account the 3d aspects of the flow. This is appropriate because the upstream flow in the tunnel is essentially laminar and at the...H2O Druck pressure transducer to measure the local cp distribution. The cp is calculated by taking the inlet total pressure from an upstream pitot
Bunnell, L. Roy; Piippo, Steven W.
1993-01-01
The objective of this educational exercise is to have students build and evaluate simple wing structures, and in doing so, learn about materials choices and lightweight construction methods. A list of equipment and supplies and the procedure for the experiment are presented.
Ali, Md. Nesar; Alam, Mahbubul; Hossain, Md. Abed; Ahmed, Md. Imteaz
2017-06-01
Flight is the main mode of locomotion used by most of the world's bird & insect species. This article discusses the mechanics of bird flight, with emphasis on the varied forms of bird's & insect's wings. The fundamentals of bird flight are similar to those of aircraft. Flying animals flap their wings to generate lift and thrust as well as to perform remarkable maneuvers with rapid accelerations and decelerations. Insects and birds provide illuminating examples of unsteady aerodynamics. Lift force is produced by the action of air flow on the wing, which is an airfoil. The airfoil is shaped such that the air provides a net upward force on the wing, while the movement of air is directed downward. Additional net lift may come from airflow around the bird's & insect's body in some species, especially during intermittent flight while the wings are folded or semi-folded. Bird's & insect's flight in nature are sub-divided into two stages. They are Unpowered Flight: Gliding and Soaring & Powered Flight: Flapping. When gliding, birds and insects obtain both a vertical and a forward force from their wings. When a bird & insect flaps, as opposed to gliding, its wings continue to develop lift as before, but the lift is rotated forward to provide thrust, which counteracts drag and increases its speed, which has the effect of also increasing lift to counteract its weight, allowing it to maintain height or to climb. Flapping flight is more complicated than flight with fixed wings because of the structural movement and the resulting unsteady fluid dynamics. Flapping involves two stages: the down-stroke, which provides the majority of the thrust, and the up-stroke, which can also (depending on the bird's & insect's wings) provide some thrust. Most kinds of bird & insect wing can be grouped into four types, with some falling between two of these types. These types of wings are elliptical wings, high speed wings, high aspect ratio wings and soaring wings with slots. Hovering is used
Restoration of the wear-resistant coatings on a GTE compressor airfoil shroud platform
Abraimov, N. V.; Geikin, V. A.; Chekalova, E. A.; Lukina, V. V.
2017-06-01
The deposition of a VT20 alloy onto the airfoil shroud platform of a compressor in an argon atmosphere and the composition, the structure, and the properties of a restored wear-resistant VK-25M coating are studied. The coating deposited onto the built-up material is found to contain (%) 3-4 C, 72-74 W, and 23-24 Co. This coating does not undergo cracking when a diamond pyramid is indented at a load of 50 kg or a diamond cone is indented at a load of 100 kg at a layer thickness of 0.15, 0.25, and 0.38 mm.
DEFF Research Database (Denmark)
van der Horst, Sander; van de Wiel, Jelmer E.; Ferreira, Carlos Simao
2016-01-01
Blades on a Vertical Axis Wind Turbine (VAWT) experience curved streamlines, caused by the rotation of the turbine. This phenomenon is known as flow curvature and has effects on the aerodynamic loading of the blades. Several authors have proposed methods to account for flow curvature, resulting...... errors remain that are intrinsic to the conformal methods used. It is shown that VAWT rotation is equivalent to an eternal pitching motion. Using this similarity, flow curvature modeling has been added to the airfoil analysis tool XFOIL. The various changes have been made in the inviscid solver, in a way...
Quantification of Airfoil Geometry-Induced Aerodynamic Uncertainties---Comparison of Approaches
Liu, Dishi
2015-04-14
Uncertainty quantification in aerodynamic simulations calls for efficient numerical methods to reduce computational cost, especially for uncertainties caused by random geometry variations which involve a large number of variables. This paper compares five methods, including quasi-Monte Carlo quadrature, polynomial chaos with coefficients determined by sparse quadrature and by point collocation, radial basis function and a gradient-enhanced version of kriging, and examines their efficiency in estimating statistics of aerodynamic performance upon random perturbation to the airfoil geometry which is parameterized by independent Gaussian variables. The results show that gradient-enhanced surrogate methods achieve better accuracy than direct integration methods with the same computational cost.
A Novel SMA-based Concept for Airfoil Structural Morphing
Barbarino, S.; Pecora, R.; Lecce, L.; Concilio, A.; Ameduri, S.; Calvi, E.
2009-08-01
The adaptive structures concept is of great interest in the aerospace field because of the several benefits which can be accomplished in the fields including noise reduction, load alleviation, weight reduction, etc., at a level in which they can be considered as compulsory in the design of future aircraft. Improvements in terms of the aerodynamic efficiency, aeroelastic behavior, stability, and manoeuvrability performance have already been proved through many international studies in the past. In the family of the Smart Materials, Shape Memory Alloys (SMA) seem to be a suitable solution for many static applications. Their high structural integrability in conjunction with actuation capabilities and a favorable performance per weight ratio, allows the development of original architectures. In this study, a morphing wing trailing edge concept is presented; morphing ability was introduced with the aim of replacing a conventional flap device. A compliant rib structure was designed, based on SMA actuators exhibiting structural potential (bearing external aerodynamic loads). Numerical results, achieved through a FE approach, are presented in terms of trailing edge induced displacement and morphed shape.
Norton, J. M.; Tari, U.; Weber, R. M.
1979-01-01
A quasi three dimensional design system and multiple-circular-arc airfoil sections were used to design a fan rotor. An axisymmetric intrablade flow field calculation modeled the shroud of an isolated splitter and radial distribution. The structural analysis indicates that the design is satisfactory for evaluation of aerodynamic performance of the fan stage in a test facility.
Axial Turbine Aerodynamic Design of Small Heavy-Duty Gas Turbines
International Nuclear Information System (INIS)
Kim, Joung Seok; Lee, Wu Sang; Ryu, Je Wook
2013-01-01
This study describes the aerodynamic design procedure for the axial turbines of a small heavy-duty gas turbine engine being developed by Docosan Heavy Industries. The design procedure mainly consists of three parts: namely, flow path design, airfoil design, and 3a performance calculation. To design the optimized flow path, through flow calculations as well as the loss estimation are widely used to evaluate the effect of geometric variables, for example, shape of meridional plane, mean radius, blades axial gap, and had angle. During the airfoil design procedure, the optimum number of blades is calculated by empirical correlations based on the in/outlet flow angles, and then 2a airfoil planar sections are designed carefully, followed by 2a B2 NS calculations. The designed planar sections are stacked along the span wise direction, leading to a 3a surfaced airfoil shape. To consider the 3a effect on turbine performance, 3a multistage Euler calculation, single row, and multistage NS calculations are performed
Topology Design of Pressure Adaptive Honeycomb for a Morphing Fowler Flap
Scheepstra, J.; Vos, R.; Barrett, R.
2011-01-01
A new method for designing a morphing Fowler flap based on pressure-adaptive honeycomb is detailed. Pressure adaptive honeycomb has been shown to be able to induce gross camber deformations in airfoil sections, such as a flap. However, due to the large amount of design variables the integration of
A parametric wing design study for a modern laminar flow wing
Koegler, J. A., Jr.
1979-01-01
The results of a parametric wing design study using a modern laminar flow airfoil designed to exhibit desirable stall characteristics while maintaining high cruise performance are presented. It was found that little is sacrificed in cruise performance when satisfying the stall margin requirements if a taper ratio of 0.65 or greater is used.
1947-06-01
airfoil was made of laminated mahogany finished with paint and 8 *- . . NM~ TN No, 1304 sanded to an aerodynamic smoothness. Fitted in only were a...cozzrespmied to a velocity of approximately 99 miles per hour and to a Remolds number of apjjroxh.nately3.68x 106. 3inge+no~nt, lift, and drag testi were
DEFF Research Database (Denmark)
Jensen, Ole B.; Pettiway, Keon
2017-01-01
by designers, planners, etc. (staging from above) and mobile subjects (staging from below). A research agenda for studying situated practices of mobility and mobilities design is outlined in three directions: foci of studies, methods and approaches, and epistemologies and frames of thinking. Jensen begins......In this chapter, Ole B. Jensen takes a situational approach to mobilities to examine how ordinary life activities are structured by technology and design. Using “staging mobilities” as a theoretical approach, Jensen considers mobilities as overlapping, actions, interactions and decisions...... with a brief description of how movement is studied within social sciences after the “mobilities turn” versus the idea of physical movement in transport geography and engineering. He then explains how “mobilities design” was derived from connections between traffic and architecture. Jensen concludes...
Computational Simulation of the Flow Past an Airfoil for an Unmanned Aerial Vehicle
Directory of Open Access Journals (Sweden)
L. Velázquez-Araque
2013-08-01
Full Text Available This paper deals with the numerical simulation of the two-dimensional, incompressible, steady air flow past a NACA 2415 airfoil and four modifications of this one. The modification of this airfoil was made in order to create a blowing outlet with the shape of a step on the suction surface. Therefore, five different locations along the cord line for this blowing outlet were analyzed. This analysis involved the aerodynamic performance which meant obtaining lift, drag and pitching moment coefficients curves as a function of the angle of attack for the situation where the engine of the aerial vehicle is turned off called the no blowing condition by means computational fluid dynamics. The RNG k-ε model is utilized to describe the turbulent flow process. The simulations were held at a Reynolds number of 105. Results allowed obtaining lift and drag forces and pitching moment coefficient and also the location of the separation and reattachment point in some cases for different angles of attack, from 0 to 16 degrees with the smallest increment of 4 degrees. Finally, numerical results were compared with results obtained from wind tunnel tests by means of an aerodynamic balance and also oil and smoke visualization techniques and found to be in very good agreement.
A data-driven decomposition approach to model aerodynamic forces on flapping airfoils
Raiola, Marco; Discetti, Stefano; Ianiro, Andrea
2017-11-01
In this work, we exploit a data-driven decomposition of experimental data from a flapping airfoil experiment with the aim of isolating the main contributions to the aerodynamic force and obtaining a phenomenological model. Experiments are carried out on a NACA 0012 airfoil in forward flight with both heaving and pitching motion. Velocity measurements of the near field are carried out with Planar PIV while force measurements are performed with a load cell. The phase-averaged velocity fields are transformed into the wing-fixed reference frame, allowing for a description of the field in a domain with fixed boundaries. The decomposition of the flow field is performed by means of the POD applied on the velocity fluctuations and then extended to the phase-averaged force data by means of the Extended POD approach. This choice is justified by the simple consideration that aerodynamic forces determine the largest contributions to the energetic balance in the flow field. Only the first 6 modes have a relevant contribution to the force. A clear relationship can be drawn between the force and the flow field modes. Moreover, the force modes are closely related (yet slightly different) to the contributions of the classic potential models in literature, allowing for their correction. This work has been supported by the Spanish MINECO under Grant TRA2013-41103-P.
Bonhaus, Daryl L.; Maddalon, Dal V.
1998-01-01
Flight-measured high Reynolds number turbulent-flow pressure distributions on a transport wing in transonic flow are compared to unstructured-grid calculations to assess the predictive ability of a three-dimensional Euler code (USM3D) coupled to an interacting boundary layer module. The two experimental pressure distributions selected for comparative analysis with the calculations are complex and turbulent but typical of an advanced technology laminar flow wing. An advancing front method (VGRID) was used to generate several tetrahedral grids for each test case. Initial calculations left considerable room for improvement in accuracy. Studies were then made of experimental errors, transition location, viscous effects, nacelle flow modeling, number and placement of spanwise boundary layer stations, and grid resolution. The most significant improvements in the accuracy of the calculations were gained by improvement of the nacelle flow model and by refinement of the computational grid. Final calculations yield results in close agreement with the experiment. Indications are that further grid refinement would produce additional improvement but would require more computer memory than is available. The appendix data compare the experimental attachment line location with calculations for different grid sizes. Good agreement is obtained between the experimental and calculated attachment line locations.
Forced Rolling Oscillation of a 65 deg-Delta Wing in Transonic Vortex-Breakdown Flow
Menzies, Margaret A.; Kandil, Osama A.; Kandil, Hamdy A.
1996-01-01
Unsteady, transonic, vortex dominated flow over a 65 deg. sharp-edged, cropped-delta wing of zero thickness undergoing forced rolling oscillations is investigated computationally. The wing angle of attack is 20 deg. and the free stream Mach number and Reynolds number are 0.85 and 3.23 x 10(exp 6), respectively. The initial condition of the flow is characterized by a transverse terminating shock which induces vortex breakdown of the leading edge vortex cores. The computational investigation uses the time accurate solution of the laminar, unsteady, compressible, full Navier-Stokes equations with the implicit, upwind, Roe flux difference splitting, finite-volume scheme. While the maximum roll amplitude is kept constant at 4.0 deg., both Reynolds number and roll frequency are varied covering three cases of forced sinusoidal rolling. First, the Reynolds number is held at 3.23 x 10(exp 6) and the wing is forced to oscillate in roll around the axis of geometric symmetry at a reduced frequency of 2(pi). Second, the Reynolds number is reduced to 0.5 x 10(exp 6) to observe the effects of added viscosity on the vortex breakdown. Third, with the Reynolds number held at 0.5 x 10(exp 6), the roll frequency is reduced to 1(pi) to complete the study.
Erdos, J. I.; Alzner, E.
1977-01-01
A numerical method of solution of the inviscid, compressible, two-dimensional unsteady flow on a blade-to-blade stream surface through a stage (rotor and stator) or a single blade row of an axial flow compressor or fan is described. A cyclic procedure has been developed for representation of adjacent blade-to-blade passages which asymptotically achieves the correct phase between all passages of a stage. A shock-capturing finite difference method is employed in the interior of the passage, and a method of characteristics technique is used at the boundaries. The blade slipstreams form two of the passage boundaries and are treated as moving contact surfaces capable of supporting jumps in entropy and tangential velocity. The Kutta condition is imposed by requiring the slipstreams to originate at the trailing edges, which are assumed to be sharp. Results are presented for several transonic fan rotors and compared with available experimental data, consisting of holographic observations of shock structure and pressure contour maps. A subcritical stator solution is also compared with results from a relaxation method. Finally, a periodic solution for a stage consisting of 44 rotor blades and 46 stator blades is discussed.
Numerical Study of Transonic Axial Flow Rotating Cascade Aerodynamics – Part 1: 2D Case
Directory of Open Access Journals (Sweden)
Irina Carmen ANDREI
2014-06-01
Full Text Available The purpose of this paper is to present a 2D study regarding the numerical simulation of flow within a transonic highly-loaded rotating cascade from an axial compressor. In order to describe an intricate flow pattern of a complex geometry and given specific conditions of cascade’s loading and operation, an appropriate accurate flow model is a must. For such purpose, the Navier-Stokes equations system was used as flow model; from the computational point of view, the mathematical support is completed by a turbulence model. A numerical comparison has been performed for different turbulence models (e.g. KE, KO, Reynolds Stress and Spallart-Allmaras models. The convergence history was monitored in order to focus on the numerical accuracy. The force vector has been reported in order to express the aerodynamics of flow within the rotating cascade at the running regime, in terms of Lift and Drag. The numerical results, expressed by plots of the most relevant flow parameters, have been compared. It comes out that the selecting of complex flow models and appropriate turbulence models, in conjunction with CFD techniques, allows to obtain the best computational accuracy of the numerical results. This paper aims to carry on a 2D study and a prospective 3D will be intended for the same architecture.
Characteristics of transonic moist air flows around butterfly valves with spontaneous condensation
Directory of Open Access Journals (Sweden)
A.B.M. Toufique Hasan
2015-06-01
Full Text Available Effects of spontaneous condensation of moist air on the shock wave dynamics around butterfly valves in transonic flows are investigated by experimental and numerical simulations. Two symmetric valve disk shapes namely- a flat rectangular plate and a mid-plane cross-section of a prototype butterfly valve have been studied in the present research. Results showed that in case with spontaneous condensation, the root mean square of pressure oscillation (induced by shock dynamics is reduced significantly with those without condensation for both shapes of the valves. Moreover, local aerodynamic moments were reduced in case with condensation which is considered to be beneficial in torque requirement in case of on/off applications of valves as flow control devices. However, total pressure loss was increased with spontaneous condensation in both the valves. Furthermore, the disk shape of a prototype butterfly valve showed better aerodynamic performances compared to flat rectangular plate profile in respect of total pressure loss and vortex shedding frequency in the wake region.
Plasma-based actuators for turbulent boundary layer control in transonic flow
Budovsky, A. D.; Polivanov, P. A.; Vishnyakov, O. I.; Sidorenko, A. A.
2017-10-01
The study is devoted to development of methods for active control of flow structure typical for the aircraft wings in transonic flow with turbulent boundary layer. The control strategy accepted in the study was based on using of the effects of plasma discharges interaction with miniature geometrical obstacles of various shapes. The conceptions were studied computationally using 3D RANS, URANS approaches. The results of the computations have shown that energy deposition can significantly change the flow pattern over the obstacles increasing their influence on the flow in boundary layer region. Namely, one of the most interesting and promising data were obtained for actuators basing on combination of vertical wedge with asymmetrical plasma discharge. The wedge considered is aligned with the local streamlines and protruding in the flow by 0.4-0.8 of local boundary layer thickness. The actuator produces negligible distortion of the flow at the absence of energy deposition. Energy deposition along the one side of the wedge results in longitudinal vortex formation in the wake of the actuator providing momentum exchange in the boundary layer. The actuator was manufactured and tested in wind tunnel experiments at Mach number 1.5 using the model of flat plate. The experimental data obtained by PIV proved the availability of the actuator.