Correlation-based Transition Modeling for External Aerodynamic Flows
NASA Astrophysics Data System (ADS)
Medida, Shivaji
Conventional turbulence models calibrated for fully turbulent boundary layers often over-predict drag and heat transfer on aerodynamic surfaces with partially laminar boundary layers. A robust correlation-based model is developed for use in Reynolds-Averaged Navier-Stokes simulations to predict laminar-to-turbulent transition onset of boundary layers on external aerodynamic surfaces. The new model is derived from an existing transition model for the two-equation k-omega Shear Stress Transport (SST) turbulence model, and is coupled with the one-equation Spalart-Allmaras (SA) turbulence model. The transition model solves two transport equations for intermittency and transition momentum thickness Reynolds number. Experimental correlations and local mean flow quantities are used in the model to account for effects of freestream turbulence level and pressure gradients on transition onset location. Transition onset is triggered by activating intermittency production using a vorticity Reynolds number criterion. In the new model, production and destruction terms of the intermittency equation are modified to improve consistency in the fully turbulent boundary layer post-transition onset, as well as ensure insensitivity to freestream eddy viscosity value specified in the SA model. In the original model, intermittency was used to control production and destruction of turbulent kinetic energy. Whereas, in the new model, only the production of eddy viscosity in SA model is controlled, and the destruction term is not altered. Unlike the original model, the new model does not use an additional correction to intermittency for separation-induced transition. Accuracy of drag predictions are improved significantly with the use of the transition model for several two-dimensional single- and multi-element airfoil cases over a wide range of Reynolds numbers. The new model is able to predict the formation of stable and long laminar separation bubbles on low-Reynolds number airfoils that
Model-based fault detection and identification with online aerodynamic model structure selection
NASA Astrophysics Data System (ADS)
Lombaerts, T.
2013-12-01
This publication describes a recursive algorithm for the approximation of time-varying nonlinear aerodynamic models by means of a joint adaptive selection of the model structure and parameter estimation. This procedure is called adaptive recursive orthogonal least squares (AROLS) and is an extension and modification of the previously developed ROLS procedure. This algorithm is particularly useful for model-based fault detection and identification (FDI) of aerospace systems. After the failure, a completely new aerodynamic model can be elaborated recursively with respect to structure as well as parameter values. The performance of the identification algorithm is demonstrated on a simulation data set.
Reduced Order Models Based on Linear and Nonlinear Aerodynamic Impulse Responses
NASA Technical Reports Server (NTRS)
Silva, Walter A.
1999-01-01
This paper discusses a method for the identification and application of reduced-order models based on linear and nonlinear aerodynamic impulse responses. The Volterra theory of nonlinear systems and an appropriate kernel identification technique are described. Insight into the nature of kernels is provided by applying the method to the nonlinear Riccati equation in a non-aerodynamic application. The method is then applied to a nonlinear aerodynamic model of an RAE 2822 supercritical airfoil undergoing plunge motions using the CFL3D Navier-Stokes flow solver with the Spalart-Allmaras turbulence model. Results demonstrate the computational efficiency of the technique.
Moving Model Test of High-Speed Train Aerodynamic Drag Based on Stagnation Pressure Measurements.
Yang, Mingzhi; Du, Juntao; Li, Zhiwei; Huang, Sha; Zhou, Dan
2017-01-01
A moving model test method based on stagnation pressure measurements is proposed to measure the train aerodynamic drag coefficient. Because the front tip of a high-speed train has a high pressure area and because a stagnation point occurs in the center of this region, the pressure of the stagnation point is equal to the dynamic pressure of the sensor tube based on the obtained train velocity. The first derivation of the train velocity is taken to calculate the acceleration of the train model ejected by the moving model system without additional power. According to Newton's second law, the aerodynamic drag coefficient can be resolved through many tests at different train speeds selected within a relatively narrow range. Comparisons are conducted with wind tunnel tests and numerical simulations, and good agreement is obtained, with differences of less than 6.1%. Therefore, the moving model test method proposed in this paper is feasible and reliable.
Moving Model Test of High-Speed Train Aerodynamic Drag Based on Stagnation Pressure Measurements
Yang, Mingzhi; Du, Juntao; Huang, Sha; Zhou, Dan
2017-01-01
A moving model test method based on stagnation pressure measurements is proposed to measure the train aerodynamic drag coefficient. Because the front tip of a high-speed train has a high pressure area and because a stagnation point occurs in the center of this region, the pressure of the stagnation point is equal to the dynamic pressure of the sensor tube based on the obtained train velocity. The first derivation of the train velocity is taken to calculate the acceleration of the train model ejected by the moving model system without additional power. According to Newton’s second law, the aerodynamic drag coefficient can be resolved through many tests at different train speeds selected within a relatively narrow range. Comparisons are conducted with wind tunnel tests and numerical simulations, and good agreement is obtained, with differences of less than 6.1%. Therefore, the moving model test method proposed in this paper is feasible and reliable. PMID:28095441
Development of Unsteady Aerodynamic State-Space Models from CFD-Based Pulse Responses
NASA Technical Reports Server (NTRS)
Silva, Walter A.; Raveh, Daniella E.
2001-01-01
A method for computing discrete-time state-space models of linearized unsteady aerodynamic behavior directly from aeroelastic CFD codes is presented. The method involves the treatment of CFD-based pulse responses as Markov parameters for use in a system identification /realization algorithm. Results are presented for the AGARD 445.6 Aeroelastic Wing with four aeroelastic modes at a Mach number of 0.96 using the EZNSS Euler/Navier-Stokes flow solver with aeroelastic capability. The System/Observer/Controller Identification Toolbox (SOCIT) algorithm, based on the Ho-Kalman realization algorithm, is used to generate 15th- and 32nd-order discrete-time state-space models of the unsteady aerodynamic response of the wing over the entire frequency range of interest.
Assessment of CFD-based Response Surface Model for Ares I Supersonic Ascent Aerodynamics
NASA Technical Reports Server (NTRS)
Hanke, Jeremy L.
2011-01-01
The Ascent Force and Moment Aerodynamic (AFMA) Databases (DBs) for the Ares I Crew Launch Vehicle (CLV) were typically based on wind tunnel (WT) data, with increments provided by computational fluid dynamics (CFD) simulations for aspects of the vehicle that could not be tested in the WT tests. During the Design Analysis Cycle 3 analysis for the outer mold line (OML) geometry designated A106, a major tunnel mishap delayed the WT test for supersonic Mach numbers (M) greater than 1.6 in the Unitary Plan Wind Tunnel at NASA Langley Research Center, and the test delay pushed the final delivery of the A106 AFMA DB back by several months. The aero team developed an interim database based entirely on the already completed CFD simulations to mitigate the impact of the delay. This CFD-based database used a response surface methodology based on radial basis functions to predict the aerodynamic coefficients for M > 1.6 based on only the CFD data from both WT and flight Reynolds number conditions. The aero team used extensive knowledge of the previous AFMA DB for the A103 OML to guide the development of the CFD-based A106 AFMA DB. This report details the development of the CFD-based A106 Supersonic AFMA DB, constructs a prediction of the database uncertainty using data available at the time of development, and assesses the overall quality of the CFD-based DB both qualitatively and quantitatively. This assessment confirms that a reasonable aerodynamic database can be constructed for launch vehicles at supersonic conditions using only CFD data if sufficient knowledge of the physics and expected behavior is available. This report also demonstrates the applicability of non-parametric response surface modeling using radial basis functions for development of aerodynamic databases that exhibit both linear and non-linear behavior throughout a large data space.
Component-based model to predict aerodynamic noise from high-speed train pantographs
NASA Astrophysics Data System (ADS)
Latorre Iglesias, E.; Thompson, D. J.; Smith, M. G.
2017-04-01
At typical speeds of modern high-speed trains the aerodynamic noise produced by the airflow over the pantograph is a significant source of noise. Although numerical models can be used to predict this they are still very computationally intensive. A semi-empirical component-based prediction model is proposed to predict the aerodynamic noise from train pantographs. The pantograph is approximated as an assembly of cylinders and bars with particular cross-sections. An empirical database is used to obtain the coefficients of the model to account for various factors: incident flow speed, diameter, cross-sectional shape, yaw angle, rounded edges, length-to-width ratio, incoming turbulence and directivity. The overall noise from the pantograph is obtained as the incoherent sum of the predicted noise from the different pantograph struts. The model is validated using available wind tunnel noise measurements of two full-size pantographs. The results show the potential of the semi-empirical model to be used as a rapid tool to predict aerodynamic noise from train pantographs.
1983-08-01
Aerodynamic Characteristics of Cruciform Missiles to High Angles of Attack Including Effects of Roll Angle and Control ...Deflections. NEAR TR 152, Nov., 1977. 2. Smith, C.A., and Nielsen, J.N.: Prediction of Aerodynamic Characteristics of Cruciform Missiles to High Angles... characteristics of body- tail and canard ( wing )- body- tail missiles . Under the same contract, the data base will be incorporated into
CFD based aerodynamic modeling to study flight dynamics of a flapping wing micro air vehicle
NASA Astrophysics Data System (ADS)
Rege, Alok Ashok
The demand for small unmanned air vehicles, commonly termed micro air vehicles or MAV's, is rapidly increasing. Driven by applications ranging from civil search-and-rescue missions to military surveillance missions, there is a rising level of interest and investment in better vehicle designs, and miniaturized components are enabling many rapid advances. The need to better understand fundamental aspects of flight for small vehicles has spawned a surge in high quality research in the area of micro air vehicles. These aircraft have a set of constraints which are, in many ways, considerably different from that of traditional aircraft and are often best addressed by a multidisciplinary approach. Fast-response non-linear controls, nano-structures, integrated propulsion and lift mechanisms, highly flexible structures, and low Reynolds aerodynamics are just a few of the important considerations which may be combined in the execution of MAV research. The main objective of this thesis is to derive a consistent nonlinear dynamic model to study the flight dynamics of micro air vehicles with a reasonably accurate representation of aerodynamic forces and moments. The research is divided into two sections. In the first section, derivation of the nonlinear dynamics of flapping wing micro air vehicles is presented. The flapping wing micro air vehicle (MAV) used in this research is modeled as a system of three rigid bodies: a body and two wings. The design is based on an insect called Drosophila Melanogaster, commonly known as fruit-fly. The mass and inertial effects of the wing on the body are neglected for the present work. The nonlinear dynamics is simulated with the aerodynamic data published in the open literature. The flapping frequency is used as the control input. Simulations are run for different cases of wing positions and the chosen parameters are studied for boundedness. Results show a qualitative inconsistency in boundedness for some cases, and demand a better
Finite Element Based Lagrangian Vortex Dynamics Model for Wind Turbine Aerodynamics
NASA Astrophysics Data System (ADS)
McWilliam, Michael K.; Crawford, Curran
2014-06-01
This paper presents a novel aerodynamic model based on Lagrangian Vortex Dynamics (LVD) formulated using a Finite Element (FE) approach. The advantage of LVD is improved fidelity over Blade Element Momentum Theory (BEMT) while being faster than Numerical Navier-Stokes Models (NNSM) in either primitive or velocity-vorticity formulations. The model improves on conventional LVD in three ways. First, the model is based on an error minimization formulation that can be solved with fast root finding algorithms. In addition to improving accuracy, this eliminates the intrinsic numerical instability of conventional relaxed wake simulations. The method has further advantages in optimization and aero-elastic simulations for two reasons. The root finding algorithm can solve the aerodynamic and structural equations simultaneously, avoiding Gauss-Seidel iteration for compatibility constraints. The second is that the formulation allows for an analytical definition for sensitivity calculations. The second improvement comes from a new discretization scheme based on an FE formulation and numerical quadrature that decouples the spatial, influencing and temporal meshes. The shape for each trailing filament uses basis functions (interpolating splines) that allow for both local polynomial order and element size refinement. A completely independent scheme distributes the influencing (vorticity) elements along the basis functions. This allows for concentrated elements in the near wake for accuracy and progressively less in the far-wake for efficiency. Finally the third improvement is the use of a far-wake model based on semi-infinite vortex cylinders where the radius and strength are related to the wake state. The error-based FE formulation allows the transition to the far wake to occur across a fixed plane.
Fourier functional analysis for unsteady aerodynamic modeling
NASA Technical Reports Server (NTRS)
Lan, C. Edward; Chin, Suei
1991-01-01
A method based on Fourier analysis is developed to analyze the force and moment data obtained in large amplitude forced oscillation tests at high angles of attack. The aerodynamic models for normal force, lift, drag, and pitching moment coefficients are built up from a set of aerodynamic responses to harmonic motions at different frequencies. Based on the aerodynamic models of harmonic data, the indicial responses are formed. The final expressions for the models involve time integrals of the indicial type advocated by Tobak and Schiff. Results from linear two- and three-dimensional unsteady aerodynamic theories as well as test data for a 70-degree delta wing are used to verify the models. It is shown that the present modeling method is accurate in producing the aerodynamic responses to harmonic motions and the ramp type motions. The model also produces correct trend for a 70-degree delta wing in harmonic motion with different mean angles-of-attack. However, the current model cannot be used to extrapolate data to higher angles-of-attack than that of the harmonic motions which form the aerodynamic model. For linear ramp motions, a special method is used to calculate the corresponding frequency and phase angle at a given time. The calculated results from modeling show a higher lift peak for linear ramp motion than for harmonic ramp motion. The current model also shows reasonably good results for the lift responses at different angles of attack.
Identification of aerodynamic models for maneuvering aircraft
NASA Technical Reports Server (NTRS)
Lan, C. Edward; Hu, C. C.
1992-01-01
A Fourier analysis method was developed to analyze harmonic forced-oscillation data at high angles of attack as functions of the angle of attack and its time rate of change. The resulting aerodynamic responses at different frequencies are used to build up the aerodynamic models involving time integrals of the indicial type. An efficient numerical method was also developed to evaluate these time integrals for arbitrary motions based on a concept of equivalent harmonic motion. The method was verified by first using results from two-dimensional and three-dimensional linear theories. The developed models for C sub L, C sub D, and C sub M based on high-alpha data for a 70 deg delta wing in harmonic motions showed accurate results in reproducing hysteresis. The aerodynamic models are further verified by comparing with test data using ramp-type motions.
NASA Technical Reports Server (NTRS)
McKann, Robert E.; Blanchard, Ulysse J.; Pearson, Albin O.
1960-01-01
The hydrodynamic and aerodynamic characteristics of a model of a multijet water-based Mach 2.0 aircraft equipped with hydrofoils have been determined. Takeoff stability and spray characteristics were very good, and sufficient excess thrust was available for takeoff in approximately 32 seconds and 4,700 feet at a gross weight of 225,000 pounds. Longitudinal and lateral stability during smooth-water landings were good. Lateral stability was good during rough-water landings, but forward location of the hydrofoils or added pitch damping was required to prevent diving. Hydrofoils were found to increase the aerodynamic lift-curve slope and to increase the aerodynamic drag coefficient in the transonic speed range, and the maximum lift-drag ratio decreased from 7.6 to 7.2 at the cruise Mach number of 0.9. The hydrofoils provided an increment of positive pitching moment over the Mach number range of the tests (0.6 to 1.42) and reduced the effective dihedral and directional stability.
Mathematical modeling of the aerodynamic characteristics in flight dynamics
NASA Technical Reports Server (NTRS)
Tobak, M.; Chapman, G. T.; Schiff, L. B.
1984-01-01
Basic concepts involved in the mathematical modeling of the aerodynamic response of an aircraft to arbitrary maneuvers are reviewed. The original formulation of an aerodynamic response in terms of nonlinear functionals is shown to be compatible with a derivation based on the use of nonlinear functional expansions. Extensions of the analysis through its natural connection with ideas from bifurcation theory are indicated.
Launch vehicle aerodynamic data base development comparison with flight data
NASA Technical Reports Server (NTRS)
Hamilton, J. T.; Wallace, R. O.; Dill, C. C.
1983-01-01
The aerodynamic development plan for the Space Shuttle integrated vehicle had three major objectives. The first objective was to support the evolution of the basic configuration by establishing aerodynamic impacts to various candidate configurations. The second objective was to provide continuing evaluation of the basic aerodynamic characteristics in order to bring about a mature data base. The third task was development of the element and component aerodynamic characteristics and distributed air loads data to support structural loads analyses. The complexity of the configurations rendered conventional analytic methods of little use and therefore required extensive wind tunnel testing of detailed complex models. However, the ground testing and analyses did not predict the aerodynamic characteristics that were extracted from the Space Shuttle flight test program. Future programs that involve the use of vehicles similar to the Space Shuttle should be concerned with the complex flow fields characteristics of these types of complex configurations.
The Crucial Role of Error Correlation for Uncertainty Modeling of CFD-Based Aerodynamics Increments
NASA Technical Reports Server (NTRS)
Hemsch, Michael J.; Walker, Eric L.
2011-01-01
The Ares I ascent aerodynamics database for Design Cycle 3 (DAC-3) was built from wind-tunnel test results and CFD solutions. The wind tunnel results were used to build the baseline response surfaces for wind-tunnel Reynolds numbers at power-off conditions. The CFD solutions were used to build increments to account for Reynolds number effects. We calculate the validation errors for the primary CFD code results at wind tunnel Reynolds number power-off conditions and would like to be able to use those errors to predict the validation errors for the CFD increments. However, the validation errors are large compared to the increments. We suggest a way forward that is consistent with common practice in wind tunnel testing which is to assume that systematic errors in the measurement process and/or the environment will subtract out when increments are calculated, thus making increments more reliable with smaller uncertainty than absolute values of the aerodynamic coefficients. A similar practice has arisen for the use of CFD to generate aerodynamic database increments. The basis of this practice is the assumption of strong correlation of the systematic errors inherent in each of the results used to generate an increment. The assumption of strong correlation is the inferential link between the observed validation uncertainties at wind-tunnel Reynolds numbers and the uncertainties to be predicted for flight. In this paper, we suggest a way to estimate the correlation coefficient and demonstrate the approach using code-to-code differences that were obtained for quality control purposes during the Ares I CFD campaign. Finally, since we can expect the increments to be relatively small compared to the baseline response surface and to be typically of the order of the baseline uncertainty, we find that it is necessary to be able to show that the correlation coefficients are close to unity to avoid overinflating the overall database uncertainty with the addition of the increments.
Experimental Facilities and Modelling for Rarefied Aerodynamics
2011-01-01
aerodynamic forces and moments that act on an object moving in the gas . The aerodynamics of rarefied gases also investigates the flow of gases in...Originally, theoretical models for rarefied gas flows were developed in the frame of the molecular kinetic theory. Thus the first self-consistent descriptions...method [7-11]. 3.0 EXPERIMENTAL FACILITIES FOR RAREFIED FLOWS 3.1 Overview Rarefied - gas (vacuum) wind tunnel is a wind tunnel operating at low pressures
NASA Technical Reports Server (NTRS)
Morelli, Eugene A.; Cunningham, Kevin; Hill, Melissa A.
2013-01-01
Flight test and modeling techniques were developed for efficiently identifying global aerodynamic models that can be used to accurately simulate stall, upset, and recovery on large transport airplanes. The techniques were developed and validated in a high-fidelity fixed-base flight simulator using a wind-tunnel aerodynamic database, realistic sensor characteristics, and a realistic flight deck representative of a large transport aircraft. Results demonstrated that aerodynamic models for stall, upset, and recovery can be identified rapidly and accurately using relatively simple piloted flight test maneuvers. Stall maneuver predictions and comparisons of identified aerodynamic models with data from the underlying simulation aerodynamic database were used to validate the techniques.
Validation and comparison of aerodynamic modelling approaches for wind turbines
NASA Astrophysics Data System (ADS)
Blondel, F.; Boisard, R.; Milekovic, M.; Ferrer, G.; Lienard, C.; Teixeira, D.
2016-09-01
The development of large capacity Floating Offshore Wind Turbines (FOWT) is an interdisciplinary challenge for the design solvers, requiring accurate modelling of both hydrodynamics, elasticity, servodynamics and aerodynamics all together. Floating platforms will induce low-frequency unsteadiness, and for large capacity turbines, the blade induced vibrations will lead to high-frequency unsteadiness. While yawed inflow conditions are still a challenge for commonly used aerodynamic methods such as the Blade Element Momentum method (BEM), the new sources of unsteadiness involved by large turbine scales and floater motions have to be tackled accurately, keeping the computational cost small enough to be compatible with design and certification purposes. In the light of this, this paper will focus on the comparison of three aerodynamic solvers based on BEM and vortex methods, on standard, yawed and unsteady inflow conditions. We will focus here on up-to-date wind tunnel experiments, such as the Unsteady Aerodynamics Experiment (UAE) database and the MexNext international project.
Xu, Gang; Liang, Xifeng; Yao, Shuanbao; Chen, Dawei
2017-01-01
Minimizing the aerodynamic drag and the lift of the train coach remains a key issue for high-speed trains. With the development of computing technology and computational fluid dynamics (CFD) in the engineering field, CFD has been successfully applied to the design process of high-speed trains. However, developing a new streamlined shape for high-speed trains with excellent aerodynamic performance requires huge computational costs. Furthermore, relationships between multiple design variables and the aerodynamic loads are seldom obtained. In the present study, the Kriging surrogate model is used to perform a multi-objective optimization of the streamlined shape of high-speed trains, where the drag and the lift of the train coach are the optimization objectives. To improve the prediction accuracy of the Kriging model, the cross-validation method is used to construct the optimal Kriging model. The optimization results show that the two objectives are efficiently optimized, indicating that the optimization strategy used in the present study can greatly improve the optimization efficiency and meet the engineering requirements. PMID:28129365
Xu, Gang; Liang, Xifeng; Yao, Shuanbao; Chen, Dawei; Li, Zhiwei
2017-01-01
Minimizing the aerodynamic drag and the lift of the train coach remains a key issue for high-speed trains. With the development of computing technology and computational fluid dynamics (CFD) in the engineering field, CFD has been successfully applied to the design process of high-speed trains. However, developing a new streamlined shape for high-speed trains with excellent aerodynamic performance requires huge computational costs. Furthermore, relationships between multiple design variables and the aerodynamic loads are seldom obtained. In the present study, the Kriging surrogate model is used to perform a multi-objective optimization of the streamlined shape of high-speed trains, where the drag and the lift of the train coach are the optimization objectives. To improve the prediction accuracy of the Kriging model, the cross-validation method is used to construct the optimal Kriging model. The optimization results show that the two objectives are efficiently optimized, indicating that the optimization strategy used in the present study can greatly improve the optimization efficiency and meet the engineering requirements.
A Generic Nonlinear Aerodynamic Model for Aircraft
NASA Technical Reports Server (NTRS)
Grauer, Jared A.; Morelli, Eugene A.
2014-01-01
A generic model of the aerodynamic coefficients was developed using wind tunnel databases for eight different aircraft and multivariate orthogonal functions. For each database and each coefficient, models were determined using polynomials expanded about the state and control variables, and an othgonalization procedure. A predicted squared-error criterion was used to automatically select the model terms. Modeling terms picked in at least half of the analyses, which totalled 45 terms, were retained to form the generic nonlinear aerodynamic (GNA) model. Least squares was then used to estimate the model parameters and associated uncertainty that best fit the GNA model to each database. Nonlinear flight simulations were used to demonstrate that the GNA model produces accurate trim solutions, local behavior (modal frequencies and damping ratios), and global dynamic behavior (91% accurate state histories and 80% accurate aerodynamic coefficient histories) under large-amplitude excitation. This compact aerodynamics model can be used to decrease on-board memory storage requirements, quickly change conceptual aircraft models, provide smooth analytical functions for control and optimization applications, and facilitate real-time parametric system identification.
Efficient Global Aerodynamic Modeling from Flight Data
NASA Technical Reports Server (NTRS)
Morelli, Eugene A.
2012-01-01
A method for identifying global aerodynamic models from flight data in an efficient manner is explained and demonstrated. A novel experiment design technique was used to obtain dynamic flight data over a range of flight conditions with a single flight maneuver. Multivariate polynomials and polynomial splines were used with orthogonalization techniques and statistical modeling metrics to synthesize global nonlinear aerodynamic models directly and completely from flight data alone. Simulation data and flight data from a subscale twin-engine jet transport aircraft were used to demonstrate the techniques. Results showed that global multivariate nonlinear aerodynamic dependencies could be accurately identified using flight data from a single maneuver. Flight-derived global aerodynamic model structures, model parameter estimates, and associated uncertainties were provided for all six nondimensional force and moment coefficients for the test aircraft. These models were combined with a propulsion model identified from engine ground test data to produce a high-fidelity nonlinear flight simulation very efficiently. Prediction testing using a multi-axis maneuver showed that the identified global model accurately predicted aircraft responses.
NASA Technical Reports Server (NTRS)
See, M. J.; Cozzolongo, J. V.
1983-01-01
A more automated process to produce wind tunnel models using existing facilities is discussed. A process was sought to more rapidly determine the aerodynamic characteristics of advanced aircraft configurations. Such aerodynamic characteristics are determined from theoretical analyses and wind tunnel tests of the configurations. Computers are used to perform the theoretical analyses, and a computer aided manufacturing system is used to fabricate the wind tunnel models. In the past a separate set of input data describing the aircraft geometry had to be generated for each process. This process establishes a common data base by enabling the computer aided manufacturing system to use, via a software interface, the geometric input data generated for the theoretical analysis. Thus, only one set of geometric data needs to be generated. Tests reveal that the process can reduce by several weeks the time needed to produce a wind tunnel model component. In addition, this process increases the similarity of the wind tunnel model to the mathematical model used by the theoretical aerodynamic analysis programs. Specifically, the wind tunnel model can be machined to within 0.008 in. of the original mathematical model. However, the software interface is highly complex and cumbersome to operate, making it unsuitable for routine use. The procurement of an independent computer aided design/computer aided manufacturing system with the capability to support both the theoretical analysis and the manufacturing tasks was recommended.
Rarefield-Flow Shuttle Aerodynamics Flight Model
NASA Technical Reports Server (NTRS)
Blanchard, Robert C.; Larman, Kevin T.; Moats, Christina D.
1994-01-01
A model of the Shuttle Orbiter rarefied-flow aerodynamic force coefficients has been derived from the ratio of flight acceleration measurements. The in-situ, low-frequency (less than 1Hz), low-level (approximately 1 x 10(exp -6) g) acceleration measurements are made during atmospheric re-entry. The experiment equipment designed and used for this task is the High Resolution Accelerometer Package (HiRAP), one of the sensor packages in the Orbiter Experiments Program. To date, 12 HiRAP re-entry mission data sets spanning a period of about 10 years have been processed. The HiRAP-derived aerodynamics model is described in detail. The model includes normal and axial hypersonic continuum coefficient equations as function of angle of attack, body-flap deflection, and elevon deflection. Normal and axial free molecule flow coefficient equations as a function of angle of attack are also presented, along with flight-derived rarefied-flow transition bridging formulae. Comparisons are made between the aerodynamics model, data from the latest Orbiter Operational Aerodynamic Design Data Book, applicable computer simulations, and wind-tunnel data.
Identification of aerodynamic models for maneuvering aircraft
NASA Technical Reports Server (NTRS)
Lan, C. Edward; Hu, C. C.
1992-01-01
The method based on Fourier functional analysis and indicial formulation for aerodynamic modeling as proposed by Chin and Lan is extensively examined and improved for the purpose of general applications to realistic airplane configurations. Improvement is made to automate the calculation of model coefficients, and to evaluate more accurately the indicial integral. Test data of large angle-of-attack ranges for two different models, a 70 deg. delta wing and an F-18 model, are used to further verify the applicability of Fourier functional analysis and validate the indicial formulation. The results show that the general expression for harmonic motions throughout a range of k is capable of accurately modeling the nonlinear responses with large phase lag except in the region where an inconsistent hysteresis behavior from one frequency to the other occurs. The results by the indicial formulation indicate that more accurate results can be obtained when the motion starts from a low angle of attack where hysteresis effect is not important.
Parameter identification and modeling of longitudinal aerodynamics
NASA Technical Reports Server (NTRS)
Aksteter, J. W.; Parks, E. K.; Bach, R. E., Jr.
1995-01-01
Using a comprehensive flight test database and a parameter identification software program produced at NASA Ames Research Center, a math model of the longitudinal aerodynamics of the Harrier aircraft was formulated. The identification program employed the equation error method using multiple linear regression to estimate the nonlinear parameters. The formulated math model structure adhered closely to aerodynamic and stability/control theory, particularly with regard to compressibility and dynamic manoeuvring. Validation was accomplished by using a three degree-of-freedom nonlinear flight simulator with pilot inputs from flight test data. The simulation models agreed quite well with the measured states. It is important to note that the flight test data used for the validation of the model was not used in the model identification.
Modeling of aircraft unsteady aerodynamic characteristics. Part 1: Postulated models
NASA Technical Reports Server (NTRS)
Klein, Vladislav; Noderer, Keith D.
1994-01-01
A short theoretical study of aircraft aerodynamic model equations with unsteady effects is presented. The aerodynamic forces and moments are expressed in terms of indicial functions or internal state variables. The first representation leads to aircraft integro-differential equations of motion; the second preserves the state-space form of the model equations. The formulations of unsteady aerodynamics is applied in two examples. The first example deals with a one-degree-of-freedom harmonic motion about one of the aircraft body axes. In the second example, the equations for longitudinal short-period motion are developed. In these examples, only linear aerodynamic terms are considered. The indicial functions are postulated as simple exponentials and the internal state variables are governed by linear, time-invariant, first-order differential equations. It is shown that both approaches to the modeling of unsteady aerodynamics lead to identical models.
Flight Test Maneuvers for Efficient Aerodynamic Modeling
NASA Technical Reports Server (NTRS)
Morelli, Eugene A.
2011-01-01
Novel flight test maneuvers for efficient aerodynamic modeling were developed and demonstrated in flight. Orthogonal optimized multi-sine inputs were applied to aircraft control surfaces to excite aircraft dynamic response in all six degrees of freedom simultaneously while keeping the aircraft close to chosen reference flight conditions. Each maneuver was designed for a specific modeling task that cannot be adequately or efficiently accomplished using conventional flight test maneuvers. All of the new maneuvers were first described and explained, then demonstrated on a subscale jet transport aircraft in flight. Real-time and post-flight modeling results obtained using equation-error parameter estimation in the frequency domain were used to show the effectiveness and efficiency of the new maneuvers, as well as the quality of the aerodynamic models that can be identified from the resultant flight data.
Identification of aerodynamic models for maneuvering aircraft
NASA Technical Reports Server (NTRS)
Chin, Suei; Lan, C. Edward
1990-01-01
Due to the requirement of increased performance and maneuverability, the flight envelope of a modern fighter is frequently extended to the high angle-of-attack regime. Vehicles maneuvering in this regime are subjected to nonlinear aerodynamic loads. The nonlinearities are due mainly to three-dimensional separated flow and concentrated vortex flow that occur at large angles of attack. Accurate prediction of these nonlinear airloads is of great importance in the analysis of a vehicle's flight motion and in the design of its flight control system. A satisfactory evaluation of the performance envelope of the aircraft may require a large number of coupled computations, one for each change in initial conditions. To avoid the disadvantage of solving the coupled flow-field equations and aircraft's motion equations, an alternate approach is to use a mathematical modeling to describe the steady and unsteady aerodynamics for the aircraft equations of motion. Aerodynamic forces and moments acting on a rapidly maneuvering aircraft are, in general, nonlinear functions of motion variables, their time rate of change, and the history of maneuvering. A numerical method was developed to analyze the nonlinear and time-dependent aerodynamic response to establish the generalized indicial function in terms of motion variables and their time rates of change.
Rarefied-flow Shuttle aerodynamics model
NASA Technical Reports Server (NTRS)
Blanchard, Robert C.; Larman, Kevin T.; Moats, Christina D.
1993-01-01
A rarefied-flow shuttle aerodynamic model spanning the hypersonic continuum to the free molecule-flow regime was formulated. The model development has evolved from the High Resolution Accelerometer Package (HiRAP) experiment conducted on the Orbiter since 1983. The complete model is described in detail. The model includes normal and axial hypersonic continuum coefficient equations as functions of angle-of-attack, body flap deflection, and elevon deflection. Normal and axial free molecule flow coefficient equations as a function of angle-of-attack are presented, along with flight derived rarefied-flow transition bridging formulae. Comparisons are made with data from the Operational Aerodynamic Design Data Book (OADDB), applicable wind-tunnel data, and recent flight data from STS-35 and STS-40. The flight-derived model aerodynamic force coefficient ratio is in good agreement with the wind-tunnel data and predicts the flight measured force coefficient ratios on STS-35 and STS-40. The model is not, however, in good agreement with the OADDB. But, the current OADDB does not predict the flight data force coefficient ratios of either STS-35 or STS-40 as accurately as the flight-derived model. Also, the OADDB differs with the wind-tunnel force coefficient ratio data.
NASA Technical Reports Server (NTRS)
Messina, Michael D.
1995-01-01
The method described in this report is intended to present an overview of a process developed to extract the forebody aerodynamic increments from flight tests. The process to determine the aerodynamic increments (rolling pitching, and yawing moments, Cl, Cm, Cn, respectively) for the forebody strake controllers added to the F/A - 18 High Alpha Research Vehicle (HARV) aircraft was developed to validate the forebody strake aerodynamic model used in simulation.
Unsteady Aerodynamic Model Tuning for Precise Flutter Prediction
NASA Technical Reports Server (NTRS)
Pak, Chan-Gi
2011-01-01
A simple method for an unsteady aerodynamic model tuning is proposed in this study. This method is based on the direct modification of the aerodynamic influence coefficient matrices. The aerostructures test wing 2 flight-test data is used to demonstrate the proposed model tuning method. The flutter speed margin computed using only the test validated structural dynamic model can be improved using the additional unsteady aerodynamic model tuning, and then the flutter speed margin requirement of 15 % in military specifications can apply towards the test validated aeroelastic model. In this study, unsteady aerodynamic model tunings are performed at two time invariant flight conditions, at Mach numbers of 0.390 and 0.456. When the Mach number for the unsteady model tuning approaches to the measured fluttering Mach number, 0.502, at the flight altitude of 9,837 ft, the estimated flutter speed is approached to the measured flutter speed at this altitude. The minimum flutter speed difference between the estimated and measured flutter speed is -.14 %.
Aerodynamic measurement techniques. [laser based diagnostic techniques
NASA Technical Reports Server (NTRS)
Hunter, W. W., Jr.
1976-01-01
Laser characteristics of intensity, monochromatic, spatial coherence, and temporal coherence were developed to advance laser based diagnostic techniques for aerodynamic related research. Two broad categories of visualization and optical measurements were considered, and three techniques received significant attention. These are holography, laser velocimetry, and Raman scattering. Examples of the quantitative laser velocimeter and Raman scattering measurements of velocity, temperature, and density indicated the potential of these nonintrusive techniques.
Development of the X-33 Aerodynamic Uncertainty Model
NASA Technical Reports Server (NTRS)
Cobleigh, Brent R.
1998-01-01
An aerodynamic uncertainty model for the X-33 single-stage-to-orbit demonstrator aircraft has been developed at NASA Dryden Flight Research Center. The model is based on comparisons of historical flight test estimates to preflight wind-tunnel and analysis code predictions of vehicle aerodynamics documented during six lifting-body aircraft and the Space Shuttle Orbiter flight programs. The lifting-body and Orbiter data were used to define an appropriate uncertainty magnitude in the subsonic and supersonic flight regions, and the Orbiter data were used to extend the database to hypersonic Mach numbers. The uncertainty data consist of increments or percentage variations in the important aerodynamic coefficients and derivatives as a function of Mach number along a nominal trajectory. The uncertainty models will be used to perform linear analysis of the X-33 flight control system and Monte Carlo mission simulation studies. Because the X-33 aerodynamic uncertainty model was developed exclusively using historical data rather than X-33 specific characteristics, the model may be useful for other lifting-body studies.
Modeling Powered Aerodynamics for the Orion Launch Abort Vehicle Aerodynamic Database
NASA Technical Reports Server (NTRS)
Chan, David T.; Walker, Eric L.; Robinson, Philip E.; Wilson, Thomas M.
2011-01-01
Modeling the aerodynamics of the Orion Launch Abort Vehicle (LAV) has presented many technical challenges to the developers of the Orion aerodynamic database. During a launch abort event, the aerodynamic environment around the LAV is very complex as multiple solid rocket plumes interact with each other and the vehicle. It is further complicated by vehicle separation events such as between the LAV and the launch vehicle stack or between the launch abort tower and the crew module. The aerodynamic database for the LAV was developed mainly from wind tunnel tests involving powered jet simulations of the rocket exhaust plumes, supported by computational fluid dynamic simulations. However, limitations in both methods have made it difficult to properly capture the aerodynamics of the LAV in experimental and numerical simulations. These limitations have also influenced decisions regarding the modeling and structure of the aerodynamic database for the LAV and led to compromises and creative solutions. Two database modeling approaches are presented in this paper (incremental aerodynamics and total aerodynamics), with examples showing strengths and weaknesses of each approach. In addition, the unique problems presented to the database developers by the large data space required for modeling a launch abort event illustrate the complexities of working with multi-dimensional data.
NASA Technical Reports Server (NTRS)
Dill, C. C.; Young, J. C.; Roberts, B. B.; Craig, M. K.; Hamilton, J. T.; Boyle, W. W.
1985-01-01
The phase B Space Shuttle systems definition studies resulted in a generic configuration consisting of a delta wing orbiter, and two solid rocket boosters (SRB) attached to an external fuel tank (ET). The initial challenge facing the aerodynamic community was aerodynamically optimizing, within limits, this configuration. As the Shuttle program developed and the sensitivities of the vehicle to aerodynamics were better understood the requirements of the aerodynamic data base grew. Adequately characterizing the vehicle to support the various design studies exploded the size of the data base to proportions that created a data modeling/management challenge for the aerodynamicist. The ascent aerodynamic data base originated primarily from wind tunnel test results. The complexity of the configuration rendered conventional analytic methods of little use. Initial wind tunnel tests provided results which included undesirable effects from model support tructure, inadequate element proximity, and inadequate plume simulation. The challenge to improve the quality of test results by determining the extent of these undesirable effects and subsequently develop testing techniques to eliminate them was imposed on the aerodynamic community. The challenges to the ascent aerodynamics community documented are unique due to the aerodynamic complexity of the Shuttle launch. Never before was such a complex vehicle aerodynamically characterized. The challenges were met with innovative engineering analyses/methodology development and wind tunnel testing techniques.
Generic Wing-Body Aerodynamics Data Base
NASA Technical Reports Server (NTRS)
Holst, Terry L.; Olsen, Thomas H.; Kwak, Dochan (Technical Monitor)
2001-01-01
The wing-body aerodynamics data base consists of a series of CFD (Computational Fluid Dynamics) simulations about a generic wing body configuration consisting of a ogive-circular-cylinder fuselage and a simple symmetric wing mid-mounted on the fuselage. Solutions have been obtained for Nonlinear Potential (P), Euler (E) and Navier-Stokes (N) solvers over a range of subsonic and transonic Mach numbers and angles of attack. In addition, each solution has been computed on a series of grids, coarse, medium and fine to permit an assessment of grid refinement errors.
System Identification of a Vortex Lattice Aerodynamic Model
NASA Technical Reports Server (NTRS)
Juang, Jer-Nan; Kholodar, Denis; Dowell, Earl H.
2001-01-01
The state-space presentation of an aerodynamic vortex model is considered from a classical and system identification perspective. Using an aerodynamic vortex model as a numerical simulator of a wing tunnel experiment, both full state and limited state data or measurements are considered. Two possible approaches for system identification are presented and modal controllability and observability are also considered. The theory then is applied to the system identification of a flow over an aerodynamic delta wing and typical results are presented.
Micro air vehicle motion tracking and aerodynamic modeling
NASA Astrophysics Data System (ADS)
Uhlig, Daniel V.
exhibited quasi-steady effects caused by small variations in the angle of attack. The quasi-steady effects, or small unsteady effects, caused variations in the aerodynamic characteristics (particularly incrementing the lift curve), and the magnitude of the influence depended on the angle-of-attack rate. In addition to nominal gliding flight, MAVs in general are capable of flying over a wide flight envelope including agile maneuvers such as perching, hovering, deep stall and maneuvering in confined spaces. From the captured motion trajectories, the aerodynamic characteristics during the numerous unsteady flights were gathered without the complexity required for unsteady wind tunnel tests. Experimental results for the MAVs show large flight envelopes that included high angles of attack (on the order of 90 deg) and high angular rates, and the aerodynamic coefficients had dynamic stall hysteresis loops and large values. From the large number of unsteady high angle-of-attack flights, an aerodynamic modeling method was developed and refined for unsteady MAV flight at high angles of attack. The method was based on a separation parameter that depended on the time history of the angle of attack and angle-of-attack rate. The separation parameter accounted for the time lag inherit in the longitudinal characteristics during dynamic maneuvers. The method was applied to three MAVs and showed general agreement with unsteady experimental results and with nominal gliding flight results. The flight tests with the MAVs indicate that modern motion tracking systems are capable of capturing the flight trajectories, and the captured trajectories can be used to determine the aerodynamic characteristics. From the captured trajectories, low Reynolds number MAV flight is explored in both nominal gliding flight and unsteady high angle-of-attack flight. Building on the experimental results, a modeling method for the longitudinal characteristics is developed that is applicable to the full flight
A workstation based simulator for teaching compressible aerodynamics
NASA Technical Reports Server (NTRS)
Benson, Thomas J.
1994-01-01
A workstation-based interactive flow simulator has been developed to aid in the teaching of undergraduate compressible aerodynamics. By solving the equations found in NACA 1135, the simulator models three basic fluids problems encountered in supersonic flow: flow past a compression corner, flow past two wedges in series, and flow past two opposed wedges. The study can vary the geometry or flow conditions through a graphical user interface and the new conditions are calculated immediately. Various graphical formats present the results of the flow calculations to the student. The simulator includes interactive questions and answers to aid in both the use of the tool and to develop an understanding of some of the complexities of compressible aerodynamics. A series of help screens make the simulator easy to learn and use.
Aerodynamics modeling of towed-cable dynamics
Kang, S.W.; Latorre, V.R.
1991-01-17
The dynamics of a cable/drogue system being towed by an orbiting aircraft has been investigated as a part of an LTWA project for the Naval Air Systems Command. We present here a status report on the tasks performed under Phase 1. We have accomplished the following tasks under Phase 1: A literature survey on the towed-cable motion problem has been conducted. While both static (steady-state) and dynamic (transient) analyses exist in the literature, no single, comprehensive analysis exists that directly addresses the present problem. However, the survey also reveals that, when judiciously applied, these past analyses can serve as useful building blocks for approaching the present problem. A numerical model that addresses several aspects of the towed-cable dynamic problem has been adapted from a Canadian underwater code for the present aerodynamic situation. This modified code, called TOWDYN, analyzes the effects of gravity, tension, aerodynamic drag, and wind. Preliminary results from this code demonstrate that the wind effects alone CAN generate the drogue oscillation behavior, i.e., the yo-yo'' phenomenon. This code also will serve as a benchmark code for checking the accuracy of a more general and complete R D'' model code. We have initiated efforts to develop a general R D model supercomputer code that also takes into account other physical factors, such as induced oscillations and bending stiffness. This general code will be able to evaluate the relative impacts of the various physical parameters, which may become important under certain conditions. This R D code will also enable development of a simpler operational code that can be used by the Naval Air personnel in the field.
Aerodynamic tailoring of the Learjet Model 60 wing
NASA Technical Reports Server (NTRS)
Chandrasekharan, Reuben M.; Hawke, Veronica M.; Hinson, Michael L.; Kennelly, Robert A., Jr.; Madson, Michael D.
1993-01-01
The wing of the Learjet Model 60 was tailored for improved aerodynamic characteristics using the TRANAIR transonic full-potential computational fluid dynamics (CFD) code. A root leading edge glove and wing tip fairing were shaped to reduce shock strength, improve cruise drag and extend the buffet limit. The aerodynamic design was validated by wind tunnel test and flight test data.
Simultaneous Excitation of Multiple-Input Multiple-Output CFD-Based Unsteady Aerodynamic Systems
NASA Technical Reports Server (NTRS)
Silva, Walter A.
2008-01-01
A significant improvement to the development of CFD-based unsteady aerodynamic reduced-order models (ROMs) is presented. This improvement involves the simultaneous excitation of the structural modes of the CFD-based unsteady aerodynamic system that enables the computation of the unsteady aerodynamic state-space model using a single CFD execution, independent of the number of structural modes. Four different types of inputs are presented that can be used for the simultaneous excitation of the structural modes. Results are presented for a flexible, supersonic semi-span configuration using the CFL3Dv6.4 code.
Simultaneous Excitation of Multiple-Input Multiple-Output CFD-Based Unsteady Aerodynamic Systems
NASA Technical Reports Server (NTRS)
Silva, Walter A.
2007-01-01
A significant improvement to the development of CFD-based unsteady aerodynamic reduced-order models (ROMs) is presented. This improvement involves the simultaneous excitation of the structural modes of the CFD-based unsteady aerodynamic system that enables the computation of the unsteady aerodynamic state-space model using a single CFD execution, independent of the number of structural modes. Four different types of inputs are presented that can be used for the simultaneous excitation of the structural modes. Results are presented for a flexible, supersonic semi-span configuration using the CFL3Dv6.4 code.
Modeling Aerodynamically Generated Sound of Helicopter Rotors
NASA Technical Reports Server (NTRS)
Brentner, Kenneth S.; Farassat, F.
2002-01-01
A great deal of progress has been made in the modeling of aerodynamically generated sound of rotors over the past decade. Although the modeling effort has focused on helicopter main rotors, the theory is generally valid for a wide range of rotor configurations. The Ffowcs Williams Hawkings (FW-H) equation has been the foundation for much of the development. The monopole and dipole source terms of the FW-H equation account for the thickness and loading noise, respectively. Bladevortex-interaction noise and broadband noise are important types of loading noise, hence much research has been directed toward the accurate modeling of these noise mechanisms. Both subsonic and supersonic quadrupole noise formulations have been developed for the prediction of high-speed impulsive noise. In an effort to eliminate the need to compute the quadrupole contribution, the FW-H equation has also been utilized on permeable surfaces surrounding all physical noise sources. Comparisons of the Kirchhoff formulation for moving surfaces with the FW-H equation have shown that the Kirchhoff formulation for moving surfaces can give erroneous results for aeroacoustic problems. Finally, significant progress has been made incorporating the rotor noise models into full vehicle noise prediction tools.
CFD Modeling of Launch Vehicle Aerodynamic Heating
NASA Technical Reports Server (NTRS)
Tashakkor, Scott B.; Canabal, Francisco; Mishtawy, Jason E.
2011-01-01
The Loci-CHEM 3.2 Computational Fluid Dynamics (CFD) code is being used to predict Ares-I launch vehicle aerodynamic heating. CFD has been used to predict both ascent and stage reentry environments and has been validated against wind tunnel tests and the Ares I-X developmental flight test. Most of the CFD predictions agreed with measurements. On regions where mismatches occurred, the CFD predictions tended to be higher than measured data. These higher predictions usually occurred in complex regions, where the CFD models (mainly turbulence) contain less accurate approximations. In some instances, the errors causing the over-predictions would cause locations downstream to be affected even though the physics were still being modeled properly by CHEM. This is easily seen when comparing to the 103-AH data. In the areas where predictions were low, higher grid resolution often brought the results closer to the data. Other disagreements are attributed to Ares I-X hardware not being present in the grid, as a result of computational resources limitations. The satisfactory predictions from CHEM provide confidence that future designs and predictions from the CFD code will provide an accurate approximation of the correct values for use in design and other applications
Aerodynamic Properties Analysis of Rapid Prototyped Models Versus Conventional Machined Models
NASA Technical Reports Server (NTRS)
Springer, A.; Cooper, K.
1998-01-01
Initial studies of the aerodynamic characteristics of proposed launch vehicles can be made more accurately if lower cost, high fidelity aerodynamic models are available for wind tunnel testing early in the design phase. This paper discusses the results of a study undertaken at NASA's Marshall Space Flight Center to determine if four rapid prototyping methods using a variety of materials are suitable for the design and manufacturing of high speed wind tunnel models in direct testing applications. It also gives an analysis of whether these materials and processes are of sufficient strength and fidelity to withstand the testing environment. In addition to test data, costs and turn-around times for the various models are given. Based on the results of this study, it can be concluded that rapid prototyping models show promise in limited direct application for preliminary aerodynamic development studies at subsonic, transonic, and supersonic speeds.
Aerodynamic Effects and Modeling of Damage to Transport Aircraft
NASA Technical Reports Server (NTRS)
Shah, Gautam H.
2008-01-01
A wind tunnel investigation was conducted to measure the aerodynamic effects of damage to lifting and stability/control surfaces of a commercial transport aircraft configuration. The modeling of such effects is necessary for the development of flight control systems to recover aircraft from adverse, damage-related loss-of-control events, as well as for the estimation of aerodynamic characteristics from flight data under such conditions. Damage in the form of partial or total loss of area was applied to the wing, horizontal tail, and vertical tail. Aerodynamic stability and control implications of damage to each surface are presented, to aid in the identification of potential boundaries in recoverable stability or control degradation. The aerodynamic modeling issues raised by the wind tunnel results are discussed, particularly the additional modeling requirements necessitated by asymmetries due to damage, and the potential benefits of such expanded modeling.
Fast-Running Aeroelastic Code Based on Unsteady Linearized Aerodynamic Solver Developed
NASA Technical Reports Server (NTRS)
Reddy, T. S. R.; Bakhle, Milind A.; Keith, T., Jr.
2003-01-01
The NASA Glenn Research Center has been developing aeroelastic analyses for turbomachines for use by NASA and industry. An aeroelastic analysis consists of a structural dynamic model, an unsteady aerodynamic model, and a procedure to couple the two models. The structural models are well developed. Hence, most of the development for the aeroelastic analysis of turbomachines has involved adapting and using unsteady aerodynamic models. Two methods are used in developing unsteady aerodynamic analysis procedures for the flutter and forced response of turbomachines: (1) the time domain method and (2) the frequency domain method. Codes based on time domain methods require considerable computational time and, hence, cannot be used during the design process. Frequency domain methods eliminate the time dependence by assuming harmonic motion and, hence, require less computational time. Early frequency domain analyses methods neglected the important physics of steady loading on the analyses for simplicity. A fast-running unsteady aerodynamic code, LINFLUX, which includes steady loading and is based on the frequency domain method, has been modified for flutter and response calculations. LINFLUX, solves unsteady linearized Euler equations for calculating the unsteady aerodynamic forces on the blades, starting from a steady nonlinear aerodynamic solution. First, we obtained a steady aerodynamic solution for a given flow condition using the nonlinear unsteady aerodynamic code TURBO. A blade vibration analysis was done to determine the frequencies and mode shapes of the vibrating blades, and an interface code was used to convert the steady aerodynamic solution to a form required by LINFLUX. A preprocessor was used to interpolate the mode shapes from the structural dynamic mesh onto the computational dynamics mesh. Then, we used LINFLUX to calculate the unsteady aerodynamic forces for a given mode, frequency, and phase angle. A postprocessor read these unsteady pressures and
Unsteady aerodynamic models for agile flight at low Reynolds numbers
NASA Astrophysics Data System (ADS)
Brunton, Steven L.
This work develops low-order models for the unsteady aerodynamic forces on a wing in response to agile maneuvers at low Reynolds number. Model performance is assessed on the basis of accuracy across a range of parameters and frequencies as well as of computational efficiency and compatibility with existing control techniques and flight dynamic models. The result is a flexible modeling procedure that yields accurate, low-dimensional, state-space models. The modeling procedures are developed and tested on direct numerical simulations of a two-dimensional flat plate airfoil in motion at low Reynolds number, Re=100, and in a wind tunnel experiment at the Illinois Institute of Technology involving a NACA 0006 airfoil pitching and plunging at Reynolds number Re=65,000. In both instances, low-order models are obtained that accurately capture the unsteady aerodynamic forces at all frequencies. These cases demonstrate the utility of the modeling procedure developed in this thesis for obtaining accurate models for different geometries and Reynolds numbers. Linear reduced-order models are constructed from either the indicial response (step response) or realistic input/output maneuvers using a flexible modeling procedure. The method is based on identifying stability derivatives and modeling the remaining dynamics with the eigensystem realization algorithm. A hierarchy of models is developed, based on linearizing the flow at various operating conditions. These models are shown to be accurate and efficient for plunging, pitching about various points, and combined pitch and plunge maneuvers, at various angle of attack and Reynolds number. Models are compared against the classical unsteady aerodynamic models of Wagner and Theodorsen over a large range of Strouhal number and reduced frequency for a baseline comparison. Additionally, state-space representations are developed for Wagner's and Theodorsen's models, making them compatible with modern control-system analysis. A number of
Bridge aerodynamics and aeroelasticity: A comparison of modeling schemes
NASA Astrophysics Data System (ADS)
Wu, Teng; Kareem, Ahsan
2013-11-01
Accurate modeling of wind-induced loads on bridge decks is critical to ensure the functionality and survivability of long-span bridges. Over the last few decades, several schemes have emerged to model bridge behavior under winds from an aerodynamic/aeroelastic perspective. A majority of these schemes rely on the quasi-steady (QS) theory. This paper systematically compares and assesses the efficacy of five analytical models available in the literature with a new model presented herein. These models include: QS theory-based model, corrected QS theory-based model, linearized QS theory-based model, semi-empirical linear model, hybrid model, and the proposed modified hybrid model. The ability of these models to capture fluid memory and nonlinear effects either individually or collectively is examined. In addition, their ability to include the effects of turbulence in the approach flow on the bridge behavior is assessed. All models are compared in a consistent manner by utilizing the time domain approach. The underlying role of each model in capturing the physics of bridge behavior under winds is highlighted and the influence of incoming turbulence and its interaction with the bridge deck is examined. A discussion is included that focuses on a number of critical parameters pivotal to the effectiveness of corresponding models.
Global Nonlinear Parametric Modeling with Application to F-16 Aerodynamics
NASA Technical Reports Server (NTRS)
Morelli, Eugene A.
1998-01-01
A global nonlinear parametric modeling technique is described and demonstrated. The technique uses multivariate orthogonal modeling functions generated from the data to determine nonlinear model structure, then expands each retained modeling function into an ordinary multivariate polynomial. The final model form is a finite multivariate power series expansion for the dependent variable in terms of the independent variables. Partial derivatives of the identified models can be used to assemble globally valid linear parameter varying models. The technique is demonstrated by identifying global nonlinear parametric models for nondimensional aerodynamic force and moment coefficients from a subsonic wind tunnel database for the F-16 fighter aircraft. Results show less than 10% difference between wind tunnel aerodynamic data and the nonlinear parameterized model for a simulated doublet maneuver at moderate angle of attack. Analysis indicated that the global nonlinear parametric models adequately captured the multivariate nonlinear aerodynamic functional dependence.
Near-wall aerodynamics of idealized model foot motion
NASA Astrophysics Data System (ADS)
Kubota, Yoshi; Hall, Joseph; Higuchi, Hiroshi; Sheth, Ritesh; Glauser, Mark; Khalifa, Ezzat
2006-11-01
The air quality is affected by amounts and types of contaminant particles suspended in the air. The particulate matter reaches the respiratory system in an indoor environment by fist becoming detached, resupended and then entrained in the human micro-environment. The resuspension phenomena from the floor occur through either a ballistic mechanism, where kinetic energy is transferred to dust particles through direct contact, or an aerodynamic mechanism, where dust particles are resuspended by the flow generated by the body. In this study we focus on the aerodynamic resuspension of particles caused by walking. The foot motion is idealized and is either towards or away from a floor. A circular disk and an elongated plate having the equivalent area to that of a human foot are used. The foot motion is driven vertically by a linear servo motor that controls the velocity, acceleration, stroke and deceleration. The model velocity is based on the real foot motion. In addition to flow visualization, flowfield measurements were conducted with PIV. In the downstroke, results show a vortex impacting the wall creating the strong wall jet. In upstroke, the vortex generated behind the idealized foot exhibits the large magnitude of velocity. Experiment is continuing with a model more closely to simulating shoe geometry as well as incorporating the real foot kinetics. The results will be compared with the numerical simulation and analytical results.
Unsteady aerodynamic modeling for arbitrary motions
NASA Technical Reports Server (NTRS)
Edwards, J. W.; Ashley, H.; Breakwell, J. V.
1977-01-01
A study is presented on the unsteady aerodynamic loads due to arbitrary motions of a thin wing and their adaptation for the calculation of response and true stability of aeroelastic modes. In an Appendix, the use of Laplace transform techniques and the generalized Theodorsen function for two-dimensional incompressible flow is reviewed. New applications of the same approach are shown also to yield airloads valid for quite general small motions. Numerical results are given for the two-dimensional supersonic case. Previously proposed approximate methods, starting from simple harmonic unsteady theory, are evaluated by comparison with exact results obtained by the present approach. The Laplace inversion integral is employed to separate the loads into 'rational' and 'nonrational' parts, of which only the former are involved in aeroelastic stability of the wing. Among other suggestions for further work, it is explained how existing aerodynamic computer programs may be adapted in a fairly straightforward fashion to deal with arbitrary transients.
Transonic limit cycle oscillation analysis using reduced order aerodynamic models
NASA Astrophysics Data System (ADS)
Dowell, E. H.; Thomas, J. P.; Hall, K. C.
2004-01-01
Limit cycle oscillations have been observed in flight operations of modern aircraft, wind tunnel experiments and mathematical models. Both fluid and structural nonlinearities are thought to contribute to these phenomena. With recent advances in reduced order aerodynamic modeling, it is now feasible to analyze limit cycle oscillations that may occur in transonic flow including the effects of structural and fluid nonlinearities. In this paper an airfoil with control surface freeplay (a common structural nonlinearity) is used to investigate transonic flutter and limit cycle oscillations. The reduced order aerodynamic model used in this paper assumes the shock motion is small and in proportion to the structural motions.
Real-Time Onboard Global Nonlinear Aerodynamic Modeling from Flight Data
NASA Technical Reports Server (NTRS)
Brandon, Jay M.; Morelli, Eugene A.
2014-01-01
Flight test and modeling techniques were developed to accurately identify global nonlinear aerodynamic models onboard an aircraft. The techniques were developed and demonstrated during piloted flight testing of an Aermacchi MB-326M Impala jet aircraft. Advanced piloting techniques and nonlinear modeling techniques based on fuzzy logic and multivariate orthogonal function methods were implemented with efficient onboard calculations and flight operations to achieve real-time maneuver monitoring and analysis, and near-real-time global nonlinear aerodynamic modeling and prediction validation testing in flight. Results demonstrated that global nonlinear aerodynamic models for a large portion of the flight envelope were identified rapidly and accurately using piloted flight test maneuvers during a single flight, with the final identified and validated models available before the aircraft landed.
Aerodynamic Models for the Low Density Supersonic Decelerator (LDSD) Test Vehicles
NASA Technical Reports Server (NTRS)
Van Norman, John W.; Dyakonov, Artem; Schoenenberger, Mark; Davis, Jody; Muppidi, Suman; Tang, Chun; Bose, Deepak; Mobley, Brandon; Clark, Ian
2016-01-01
An overview of aerodynamic models for the Low Density Supersonic Decelerator (LDSD) Supersonic Flight Dynamics Test (SFDT) campaign test vehicle is presented, with comparisons to reconstructed flight data and discussion of model updates. The SFDT campaign objective is to test Supersonic Inflatable Aerodynamic Decelerator (SIAD) and large supersonic parachute technologies at high altitude Earth conditions relevant to entry, descent, and landing (EDL) at Mars. Nominal SIAD test conditions are attained by lifting a test vehicle (TV) to 36 km altitude with a helium balloon, then accelerating the TV to Mach 4 and 53 km altitude with a solid rocket motor. Test flights conducted in June of 2014 (SFDT-1) and 2015 (SFDT-2) each successfully delivered a 6 meter diameter decelerator (SIAD-R) to test conditions and several seconds of flight, and were successful in demonstrating the SFDT flight system concept and SIAD-R technology. Aerodynamic models and uncertainties developed for the SFDT campaign are presented, including the methods used to generate them and their implementation within an aerodynamic database (ADB) routine for flight simulations. Pre- and post-flight aerodynamic models are compared against reconstructed flight data and model changes based upon knowledge gained from the flights are discussed. The pre-flight powered phase model is shown to have a significant contribution to off-nominal SFDT trajectory lofting, while coast and SIAD phase models behaved much as predicted.
Prediction and Validation of Mars Pathfinder Hypersonic Aerodynamic Data Base
NASA Technical Reports Server (NTRS)
Gnoffo, Peter A.; Braun, Robert D.; Weilmuenster, K. James; Mitcheltree, Robert A.; Engelund, Walter C.; Powell, Richard W.
1998-01-01
Postflight analysis of the Mars Pathfinder hypersonic, continuum aerodynamic data base is presented. Measured data include accelerations along the body axis and axis normal directions. Comparisons of preflight simulation and measurements show good agreement. The prediction of two static instabilities associated with movement of the sonic line from the shoulder to the nose and back was confirmed by measured normal accelerations. Reconstruction of atmospheric density during entry has an uncertainty directly proportional to the uncertainty in the predicted axial coefficient. The sensitivity of the moment coefficient to freestream density, kinetic models and center-of-gravity location are examined to provide additional consistency checks of the simulation with flight data. The atmospheric density as derived from axial coefficient and measured axial accelerations falls within the range required for sonic line shift and static stability transition as independently determined from normal accelerations.
A Rapid Aerodynamic Design Procedure Based on Artificial Neural Networks
NASA Technical Reports Server (NTRS)
Rai, Man Mohan
2001-01-01
An aerodynamic design procedure that uses neural networks to model the functional behavior of the objective function in design space has been developed. This method incorporates several improvements to an earlier method that employed a strategy called parameter-based partitioning of the design space in order to reduce the computational costs associated with design optimization. As with the earlier method, the current method uses a sequence of response surfaces to traverse the design space in search of the optimal solution. The new method yields significant reductions in computational costs by using composite response surfaces with better generalization capabilities and by exploiting synergies between the optimization method and the simulation codes used to generate the training data. These reductions in design optimization costs are demonstrated for a turbine airfoil design study where a generic shape is evolved into an optimal airfoil.
Development of Unsteady Aerodynamic and Aeroelastic Reduced-Order Models Using the FUN3D Code
NASA Technical Reports Server (NTRS)
Silva, Walter A.; Vatsa, Veer N.; Biedron, Robert T.
2009-01-01
Recent significant improvements to the development of CFD-based unsteady aerodynamic reduced-order models (ROMs) are implemented into the FUN3D unstructured flow solver. These improvements include the simultaneous excitation of the structural modes of the CFD-based unsteady aerodynamic system via a single CFD solution, minimization of the error between the full CFD and the ROM unsteady aero- dynamic solution, and computation of a root locus plot of the aeroelastic ROM. Results are presented for a viscous version of the two-dimensional Benchmark Active Controls Technology (BACT) model and an inviscid version of the AGARD 445.6 aeroelastic wing using the FUN3D code.
Evaluation of thermographic phosphor technology for aerodynamic model testing
Cates, M.R.; Tobin, K.W.; Smith, D.B.
1990-08-01
The goal for this project was to perform technology evaluations applicable to the development of higher-precision, higher-temperature aerodynamic model testing at Arnold Engineering Development Center (AEDC) in Tullahmoa, Tennessee. With the advent of new programs for design of aerospace craft that fly at higher speeds and altitudes, requirements for detailed understanding of high-temperature materials become very important. Model testing is a natural and critical part of the development of these new initiatives. The well-established thermographic phosphor techniques of the Applied Technology Division at Oak Ridge National Laboratory are highly desirable for diagnostic evaluation of materials and aerodynamic shapes as studied in model tests. Combining this state-of-the-art thermographic technique with modern, higher-temperature models will greatly improve the practicability of tests for the advanced aerospace vehicles and will provide higher precision diagnostic information for quantitative evaluation of these tests. The wavelength ratio method for measuring surface temperatures of aerodynamic models was demonstrated in measurements made for this project. In particular, it was shown that the appropriate phosphors could be selected for the temperature range up to {approximately}700 {degree}F or higher and emission line ratios of sufficient sensitivity to measure temperature with 1% precision or better. Further, it was demonstrated that two-dimensional image- processing methods, using standard hardware, can be successfully applied to surface thermography of aerodynamic models for AEDC applications.
ON AERODYNAMIC AND BOUNDARY LAYER RESISTANCES WITHIN DRY DEPOSITION MODELS
There have been many empirical parameterizations for the aerodynamic and boundary layer resistances proposed in the literature, e.g. those of the Meyers Multi-Layer Deposition Model (MLM) used with the nation-wide dry deposition network. Many include arbitrary constants or par...
NASA Technical Reports Server (NTRS)
Friedmann, P. P.; Venkatesan, C.
1985-01-01
The aeromechanical stability of a helicopter in ground resonance was analyzed, by incorporating five different aerodynamic models in the coupled rotor/fuselage analysis. The sensitivity of the results to changes in aerodynamic modelling was carefully examined. The theoretical results were compared with experimental data and useful conclusions are drawn regarding the role of aerodynamic modeling on this aeromechanical stability problem. The aerodynamic model which provided the best all around correlation with the experimental data was identified.
Modelling of natural and bypass transition in aerodynamics
NASA Astrophysics Data System (ADS)
Fürst, Jiří; Straka, Petr; Příhoda, Jaromír
2014-03-01
Most of transition models are proposed for modelling of the bypass transition common in the internal aerodynamics especially in turbomachinery where free stream turbulence is the dominant parameter affecting the transition onset. Free-stream turbulence level in the external aerodynamics is usually noticeably lower and so the natural transition often occurs in flows around airfoils. The transition model with the algebraic equation for the intermittency coefficient proposed originally by Straka and Příhoda [3] for the bypass transition was modified for modelling of the transition at low free-stream turbulence. The modification is carried out using experimental data of Schubauer and Skramstad [18]. Further, the three-equation k-kL-ω model proposed by Walters and Cokljat [10] was used for the modelling of the transition at low free-stream turbulence. Both models were tested by means of the incompressible flow around airfoils at moderate and very low free-stream turbulence.
A CFD-informed quasi-steady model of flapping wing aerodynamics
Nakata, Toshiyuki; Liu, Hao; Bomphrey, Richard J.
2016-01-01
Aerodynamic performance and agility during flapping flight are determined by the combination of wing shape and kinematics. The degree of morphological and kinematic optimisation is unknown and depends upon a large parameter space. Aimed at providing an accurate and computationally inexpensive modelling tool for flapping-wing aerodynamics, we propose a novel CFD (computational fluid dynamics)-informed quasi-steady model (CIQSM), which assumes that the aerodynamic forces on a flapping wing can be decomposed into the quasi-steady forces and parameterised based on CFD results. Using least-squares fitting, we determine a set of proportional coefficients for the quasi-steady model relating wing kinematics to instantaneous aerodynamic force and torque; we calculate power with the product of quasi-steady torques and angular velocity. With the quasi-steady model fully and independently parameterised on the basis of high-fidelity CFD modelling, it is capable of predicting flapping-wing aerodynamic forces and power more accurately than the conventional blade element model (BEM) does. The improvement can be attributed to, for instance, taking into account the effects of the induced downwash and the wing tip vortex on the force generation and power consumption. Our model is validated by comparing the aerodynamics of a CFD model and the present quasi-steady model using the example case of a hovering hawkmoth. It demonstrates that the CIQSM outperforms the conventional BEM while remaining computationally cheap, and hence can be an effective tool for revealing the mechanisms of optimization and control of kinematics and morphology in flapping-wing flight for both bio-flyers and unmanned air systems. PMID:27346891
A CFD-informed quasi-steady model of flapping wing aerodynamics.
Nakata, Toshiyuki; Liu, Hao; Bomphrey, Richard J
2015-11-01
Aerodynamic performance and agility during flapping flight are determined by the combination of wing shape and kinematics. The degree of morphological and kinematic optimisation is unknown and depends upon a large parameter space. Aimed at providing an accurate and computationally inexpensive modelling tool for flapping-wing aerodynamics, we propose a novel CFD (computational fluid dynamics)-informed quasi-steady model (CIQSM), which assumes that the aerodynamic forces on a flapping wing can be decomposed into the quasi-steady forces and parameterised based on CFD results. Using least-squares fitting, we determine a set of proportional coefficients for the quasi-steady model relating wing kinematics to instantaneous aerodynamic force and torque; we calculate power with the product of quasi-steady torques and angular velocity. With the quasi-steady model fully and independently parameterised on the basis of high-fidelity CFD modelling, it is capable of predicting flapping-wing aerodynamic forces and power more accurately than the conventional blade element model (BEM) does. The improvement can be attributed to, for instance, taking into account the effects of the induced downwash and the wing tip vortex on the force generation and power consumption. Our model is validated by comparing the aerodynamics of a CFD model and the present quasi-steady model using the example case of a hovering hawkmoth. It demonstrates that the CIQSM outperforms the conventional BEM while remaining computationally cheap, and hence can be an effective tool for revealing the mechanisms of optimization and control of kinematics and morphology in flapping-wing flight for both bio-flyers and unmanned air systems.
Estimation of Unsteady Aerodynamic Models from Dynamic Wind Tunnel Data
NASA Technical Reports Server (NTRS)
Murphy, Patrick; Klein, Vladislav
2011-01-01
Demanding aerodynamic modelling requirements for military and civilian aircraft have motivated researchers to improve computational and experimental techniques and to pursue closer collaboration in these areas. Model identification and validation techniques are key components for this research. This paper presents mathematical model structures and identification techniques that have been used successfully to model more general aerodynamic behaviours in single-degree-of-freedom dynamic testing. Model parameters, characterizing aerodynamic properties, are estimated using linear and nonlinear regression methods in both time and frequency domains. Steps in identification including model structure determination, parameter estimation, and model validation, are addressed in this paper with examples using data from one-degree-of-freedom dynamic wind tunnel and water tunnel experiments. These techniques offer a methodology for expanding the utility of computational methods in application to flight dynamics, stability, and control problems. Since flight test is not always an option for early model validation, time history comparisons are commonly made between computational and experimental results and model adequacy is inferred by corroborating results. An extension is offered to this conventional approach where more general model parameter estimates and their standard errors are compared.
Modeling and simulation of coaxial helicopter rotor aerodynamics
NASA Astrophysics Data System (ADS)
Gecgel, Murat
A framework is developed for the computational fluid dynamics (CFD) analyses of a series of helicopter rotor flowfields in hover and in forward flight. The methodology is based on the unsteady solutions of the three-dimensional, compressible Navier-Stokes equations recast in a rotating frame of reference. The simulations are carried out by solving the developed mathematical model on hybrid meshes that aim to optimally exploit the benefits of both the structured and the unstructured grids around complex configurations. The computer code is prepared for parallel processing with distributed memory utilization in order to significantly reduce the computational time and the memory requirements. The developed model and the simulation methodology are validated for single-rotor-in-hover flowfields by comparing the present results with the published experimental data. The predictive merit of different turbulence models for complex helicopter aerodynamics are tested extensively. All but the kappa-o and LES results demonstrate acceptable agreement with the experimental data. It was deemed best to use the one-equation Spalart-Allmaras turbulence model for the subsequent rotor flowfield computations. First, the flowfield around a single rotor in forward flight is simulated. These time---accurate computations help to analyze an adverse effect of increasing the forward flight speed. A dissymmetry of the lift on the advancing and the retreating blades is observed for six different advance ratios. Since the coaxial rotor is proposed to mitigate the dissymmetry, it is selected as the next logical step of the present investigation. The time---accurate simulations are successfully obtained for the flowfields generated by first a hovering then a forward-flying coaxial rotor. The results for the coaxial rotor in forward flight verify the aerodynamic balance proposed by the previously published advancing blade concept. The final set of analyses aims to investigate if the gap between the
Moniuszko, Justyna; Maryniak, Jerzy; Ladyżyńska-Kozdraś, Edyta
2010-01-01
Based on a model of a parachute jumper, for various body configurations in a sitting position, tests were carried out in an aerodynamic tunnel. Aerodynamic characteristics and dimensionless aerodynamic forces' coefficients were calculated. The tests were carried out for various configurations of the jumper's body. A universal mathematical model of a parachute jumper's body was prepared, thus enabling the analysis of the jumper's movement with a closed parachute in any position. In order to build the model, a digitized model of a jumper allowing changing the body configuration, making appropriate changes of the moment of inertia, distribution of the center of mass and the aerodynamic characteristics was adopted. Dynamic movement equations were derived for a jumper in a relative reference system. The mathematical model was formulated for a jumper with a variable body configuration during the flight, which can be realized through a change of the position and the speed of the parachute jumper's limbs. The model allows analyzing the motion of the jumper with a closed parachute. It is an important jump phase during an assault with delayed parachute opening in sports type jumping, e.g., Skydiving and in emergency jumps from higher altitudes for the parachute's opening to be safe.
Aerodynamic experimentation with ducted models as applied to hypersonic air-breathing vehicles
NASA Astrophysics Data System (ADS)
Goon'ko, Yu. P.
A methodology of experimentation in high supersonic wind tunnels for studying aerodynamic characteristics of hypersonic flying vehicles powered by air-breathing engines is discussed. Investigations of such total aerodynamic forces as drag, lift and pitching moment at testing the models are implicit when the air flow through the model ducts is accomplished so that to provide the simulation of the external flow around the airplane and flow over the inlets, but the operating engines and, hence, the exhaust jets are not modeled. The methods used for testing such models are based on the measurement of duct stream parameters alongside with the balance measurement of aerodynamic forces acting on the models. In the tests, aerometric tools are used such as narrow metering nozzles (plugs), pitot and static pressure probes, stagnation temperature probes and pressure orifices in walls of the model duct. The aerometric data serve to determine the flow rate and momentum of the stream at the duct exit. The internal non-simulated forces of the model ducts are also determined using the conservation equations for energy, mass flow and momentum, and these forces are eliminated from the aerodynamic test results. The techniques of the said model testing have been well developed as applied to supersonic aircraft, however their application for hypersonic vehicles whose models are tested at high supersonic speeds, Mach number M∞>4, implies some specific features. In the present paper, the results of experimental and theoretical study of these features are discussed. Some experimental data on aerodynamics of hypersonic aircraft models received in methodological tests are also presented. The tunnel experiments have been carried out in the Mach number range M∞=2-6.
Determining Aerodynamic Loads Based on Optical Deformation Measurements
NASA Technical Reports Server (NTRS)
Liu, Tianshu; Barrows, D. A.; Burner, A. W.; Rhew, R. D.
2001-01-01
This paper describes a videogrammetric technique for determining aerodynamic loads based on optical elastic deformation measurements. The data reduction methods are developed to extract the normal force and pitching moment from beam deformation data. The axial force is obtained by measuring the axial translational motion of a movable shaft in a spring/bearing device. Proof-of-concept calibration experiments are conducted to assess the accuracy of this optical technique.
Determining Aerodynamic Loads Based on Optical Deformation Measurements
NASA Technical Reports Server (NTRS)
Liu, Tianshu; Barrows, D. A.; Burner, A. W.; Rhew, R. D.
2001-01-01
This paper describes a videogram metric technique for determining aerodynamic loads based on optical elastic deformation measurements. The data reduction methods are developed to extract the normal force and pitching moment from beam deformation data. The axial force is obtained by measuring the axial translational motion of a movable shaft in a spring/bearing device. Proof-of-concept calibration experiments are conducted to assess the accuracy of this optical technique.
Preliminary subsonic aerodynamic model for simulation studies of the HL-20 lifting body
NASA Technical Reports Server (NTRS)
Jackson, E. Bruce; Cruz, Christopher I.
1992-01-01
A nonlinear, six-degree-of-freedom aerodynamic model for an early version of the HL-20 lifting body is described and compared with wind tunnel data upon which it is based. Polynomial functions describing most of the aerodynamic parameters are given and tables of these functions are presented. Techniques used to arrive at these functions are described. Basic aerodynamic coefficients were modeled as functions of angles of attack and sideslip. Vehicle lateral symmetry was assumed. Compressibility (Mach) effects were ignored. Control-surface effectiveness was assumed to vary linearly with angle of deflection and was assumed to be invariant with the angle of sideslip. Dynamic derivatives were obtained from predictive aerodynamic codes. Landing-gear and ground effects were scaled from Space Shuttle data. The model described is provided to support pilot-in-the-loop simulation studies of the HL-20. By providing the data in tabular format, the model is suitable for the data interpolation architecture of many existing engineering simulation facilities. Because of the preliminary nature of the data, however, this model is not recommended for study of the absolute performance of the HL-20.
Aerodynamic Shape Optimization Based on Free-form Deformation
NASA Technical Reports Server (NTRS)
Samareh, Jamshid A.
2004-01-01
This paper presents a free-form deformation technique suitable for aerodynamic shape optimization. Because the proposed technique is independent of grid topology, we can treat structured and unstructured computational fluid dynamics grids in the same manner. The proposed technique is an alternative shape parameterization technique to a trivariate volume technique. It retains the flexibility and freedom of trivariate volumes for CFD shape optimization, but it uses a bivariate surface representation. This reduces the number of design variables by an order of magnitude, and it provides much better control for surface shape changes. The proposed technique is simple, compact, and efficient. The analytical sensitivity derivatives are independent of the design variables and are easily computed for use in a gradient-based optimization. The paper includes the complete formulation and aerodynamics shape optimization results.
System Dynamic Analysis of a Wind Tunnel Model with Applications to Improve Aerodynamic Data Quality
NASA Technical Reports Server (NTRS)
Buehrle, Ralph David
1997-01-01
The research investigates the effect of wind tunnel model system dynamics on measured aerodynamic data. During wind tunnel tests designed to obtain lift and drag data, the required aerodynamic measurements are the steady-state balance forces and moments, pressures, and model attitude. However, the wind tunnel model system can be subjected to unsteady aerodynamic and inertial loads which result in oscillatory translations and angular rotations. The steady-state force balance and inertial model attitude measurements are obtained by filtering and averaging data taken during conditions of high model vibrations. The main goals of this research are to characterize the effects of model system dynamics on the measured steady-state aerodynamic data and develop a correction technique to compensate for dynamically induced errors. Equations of motion are formulated for the dynamic response of the model system subjected to arbitrary aerodynamic and inertial inputs. The resulting modal model is examined to study the effects of the model system dynamic response on the aerodynamic data. In particular, the equations of motion are used to describe the effect of dynamics on the inertial model attitude, or angle of attack, measurement system that is used routinely at the NASA Langley Research Center and other wind tunnel facilities throughout the world. This activity was prompted by the inertial model attitude sensor response observed during high levels of model vibration while testing in the National Transonic Facility at the NASA Langley Research Center. The inertial attitude sensor cannot distinguish between the gravitational acceleration and centrifugal accelerations associated with wind tunnel model system vibration, which results in a model attitude measurement bias error. Bias errors over an order of magnitude greater than the required device accuracy were found in the inertial model attitude measurements during dynamic testing of two model systems. Based on a theoretical modal
Computational modeling of aerodynamic characteristics in sprayed and spiraled precalciner
NASA Astrophysics Data System (ADS)
Li, Xiangguo; Ma, Baoguo; Hu, Zhenwu
2008-08-01
Based on the structural and work characteristics of a spiraled and sprayed precalciner, the RNG k- ɛ model and the SIMPLE method were used to simulate the aerodynamic characteristics in a sprayed and spiraled precalciner. The simulation results demonstrate that the flow area of airflow was increased abruptly due to the reduced part of the bottom of precalciners, which attributed to a sprayed effect. With the mix of the tertiary air with the swirl flow and secondary air, a high-speed zone was formed in the opposite side of the inlet of tertiary air, in which the highest speed was 32.97 m/s. Moreover, the inlet of raw meal designed in the high-speed zone can be propitious to the decentralization of the raw meal. A back-flow zone was formed near the side of the inlet of tertiary air, in which the velocity was negative. From the analysis of the results, the flow field of the precalciner is composed of a sprayed zone, a high-speed zone, a back-flow zone and cylinder zone; moreover, the simulation results agree with those of the engineering compared to the in situ results. The results also showed that the CFD method can be used to give the basis for optimizing the geometrical design and flow parameters of a precalciner.
NASA Technical Reports Server (NTRS)
Nissim, Eli
1990-01-01
The aerodynamic energy method is used to synthesize control laws for NASA's drone for aerodynamic and structural testing-aerodynamic research wing 1 (DAST-ARW1) mathematical model. The performance of these control laws in terms of closed-loop flutter dynamic pressure, control surface activity, and robustness is compared with other control laws that relate to the same model. A control law synthesis technique that makes use of the return difference singular values is developed. It is based on the aerodynamic energy approach and is shown to yield results that are superior to those results given in the literature and are based on optimal control theory. Nyquist plots are presented, together with a short discussion regarding the relative merits of the minimum singular value as a measure of robustness as compared with the more traditional measure involving phase and gain margins.
NASA Technical Reports Server (NTRS)
Nissim, E.
1989-01-01
The aerodynamic energy method is used in this paper to synthesize control laws for NASA's Drone for Aerodynamic and Structural Testing-Aerodynamic Research Wing 1 (DAST-ARW1) mathematical model. The performance of these control laws in terms of closed-loop flutter dynamic pressure, control surface activity, and robustness is compared against other control laws that appear in the literature and relate to the same model. A control law synthesis technique that makes use of the return difference singular values is developed in this paper. it is based on the aerodynamic energy approach and is shown to yield results superior to those given in the literature and based on optimal control theory. Nyquist plots are presented together with a short discussion regarding the relative merits of the minimum singular value as a measure of robustness, compared with the more traditional measure of robustness involving phase and gain margins.
Modeling of turbulent separated flows for aerodynamic applications
NASA Technical Reports Server (NTRS)
Marvin, J. G.
1983-01-01
Steady, high speed, compressible separated flows modeled through numerical simulations resulting from solutions of the mass-averaged Navier-Stokes equations are reviewed. Emphasis is placed on benchmark flows that represent simplified (but realistic) aerodynamic phenomena. These include impinging shock waves, compression corners, glancing shock waves, trailing edge regions, and supersonic high angle of attack flows. A critical assessment of modeling capabilities is provided by comparing the numerical simulations with experiment. The importance of combining experiment, numerical algorithm, grid, and turbulence model to effectively develop this potentially powerful simulation technique is stressed.
Analytical aerodynamic model of a high alpha research vehicle wind-tunnel model
NASA Technical Reports Server (NTRS)
Cao, Jichang; Garrett, Frederick, Jr.; Hoffman, Eric; Stalford, Harold
1990-01-01
A 6 DOF analytical aerodynamic model of a high alpha research vehicle is derived. The derivation is based on wind-tunnel model data valid in the altitude-Mach flight envelope centered at 15,000 ft altitude and 0.6 Mach number with Mach range between 0.3 and 0.9. The analytical models of the aerodynamics coefficients are nonlinear functions of alpha with all control variable and other states fixed. Interpolation is required between the parameterized nonlinear functions. The lift and pitching moment coefficients have unsteady flow parts due to the time range of change of angle-of-attack (alpha dot). The analytical models are plotted and compared with their corresponding wind-tunnel data. Piloted simulated maneuvers of the wind-tunnel model are used to evaluate the analytical model. The maneuvers considered are pitch-ups, 360 degree loaded and unloaded rolls, turn reversals, split S's, and level turns. The evaluation finds that (1) the analytical model is a good representation at Mach 0.6, (2) the longitudinal part is good for the Mach range 0.3 to 0.9, and (3) the lateral part is good for Mach numbers between 0.6 and 0.9. The computer simulations show that the storage requirement of the analytical model is about one tenth that of the wind-tunnel model and it runs twice as fast.
Aerodynamics of a Gulfstream G550 Nose Landing Gear Model
NASA Technical Reports Server (NTRS)
Neuhart, Dan H.; Khorrami, Mehdi R.; Choudhari, Meelan M.
2009-01-01
In this paper we discuss detailed steady and unsteady aerodynamic measurements of a Gulfstream G550 nose landing gear model. The quarter-scale, high-fidelity model includes part of the lower fuselage and the gear cavity. The full model configuration allowed for removal of various gear components (e.g. light cluster, steering mechanism, hydraulic lines, etc.) in order to document their effects on the local flow field. The measurements were conducted at a Reynolds number of 7.3 x 10(exp 4) based on the shock strut (piston) diameter and a freestream Mach number of 0.166. Additional data were also collected at lower Mach numbers of 0.12 and 0.145 and correspondingly lower Reynolds numbers. The boundary layer on the piston was tripped to enable turbulent flow separation, so as to better mimic the conditions encountered during flight. Steady surface pressures were gathered from an extensive number of static ports on the wheels, door, fuselage, and within the gear cavity. To better understand the resultant flow interactions between gear components, surface pressure fluctuations were collected via sixteen dynamic pressure sensors strategically placed on various subcomponents of the gear. Fifteen of the transducers were flush mounted on the gear surface at fixed locations, while the remaining one was a mobile transducer that could be placed at numerous varying locations. The measured surface pressure spectra are mainly broadband in nature, lacking any local peaks associated with coherent vortex shedding. This finding is in agreement with off-surface flow measurements using PIV that revealed the flow field to be a collection of separated shear layers without any dominant vortex shedding processes.
NASA Technical Reports Server (NTRS)
Klein, Vladislav; Noderer, Keith D.
1995-01-01
Aerodynamic equations with unsteady effects were formulated for an aircraft in one-degree-of-freedom, small-amplitude, harmonic motion. These equations were used as a model for aerodynamic parameter estimation from wind tunnel oscillatory data. The estimation algorithm was based on nonlinear least squares and was applied in three examples to the oscillatory data in pitch and roll of 70 deg triangular wing and an X-31 model, and in-sideslip oscillatory data of the High Incidence Research Model 2 (HIRM 2). All three examples indicated that a model using a simple indicial function can explain unsteady effects observed in measured data. The accuracy of the estimated parameters and model verification were strongly influenced by the number of data points with respect to the number of unknown parameters.
NASA Technical Reports Server (NTRS)
Gangwani, S. T.
1985-01-01
A reliable rotor aeroelastic analysis operational that correctly predicts the vibration levels for a helicopter is utilized to test various unsteady aerodynamics models with the objective of improving the correlation between test and theory. This analysis called Rotor Aeroelastic Vibration (RAVIB) computer program is based on a frequency domain forced response analysis which utilizes the transfer matrix techniques to model helicopter/rotor dynamic systems of varying degrees of complexity. The results for the AH-1G helicopter rotor were compared with the flight test data during high speed operation and they indicated a reasonably good correlation for the beamwise and chordwise blade bending moments, but for torsional moments the correlation was poor. As a result, a new aerodynamics model based on unstalled synthesized data derived from the large amplitude oscillating airfoil experiments was developed and tested.
Exploring bird aerodynamics using radio-controlled models.
Hoey, Robert G
2010-12-01
A series of radio-controlled glider models was constructed by duplicating the aerodynamic shape of soaring birds (raven, turkey vulture, seagull and pelican). Controlled tests were conducted to determine the level of longitudinal and lateral-directional static stability, and to identify the characteristics that allowed flight without a vertical tail. The use of tail-tilt for controlling small bank-angle changes, as observed in soaring birds, was verified. Subsequent tests, using wing-tip ailerons, inferred that birds use a three-dimensional flow pattern around the wing tip (wing tip vortices) to control adverse yaw and to create a small amount of forward thrust in gliding flight.
NASA Technical Reports Server (NTRS)
Tang, Chun; Muppidi, Suman; Bose, Deepak; Van Norman, John W.; Tanimoto, Rebekah; Clark, Ian
2015-01-01
NASA's Low Density Supersonic Decelerator Program is developing new technologies that will enable the landing of heavier payloads in low density environments, such as Mars. A recent flight experiment conducted high above the Hawaiian Islands has demonstrated the performance of several decelerator technologies. In particular, the deployment of the Robotic class Supersonic Inflatable Aerodynamic Decelerator (SIAD-R) was highly successful, and valuable data were collected during the test flight. This paper outlines the Computational Fluid Dynamics (CFD) analysis used to estimate the aerodynamic and aerothermal characteristics of the SIAD-R. Pre-flight and post-flight predictions are compared with the flight data, and a very good agreement in aerodynamic force and moment coefficients is observed between the CFD solutions and the reconstructed flight data.
Quasi steady-state aerodynamic model development for race vehicle simulations
NASA Astrophysics Data System (ADS)
Mohrfeld-Halterman, J. A.; Uddin, M.
2016-01-01
Presented in this paper is a procedure to develop a high fidelity quasi steady-state aerodynamic model for use in race car vehicle dynamic simulations. Developed to fit quasi steady-state wind tunnel data, the aerodynamic model is regressed against three independent variables: front ground clearance, rear ride height, and yaw angle. An initial dual range model is presented and then further refined to reduce the model complexity while maintaining a high level of predictive accuracy. The model complexity reduction decreases the required amount of wind tunnel data thereby reducing wind tunnel testing time and cost. The quasi steady-state aerodynamic model for the pitch moment degree of freedom is systematically developed in this paper. This same procedure can be extended to the other five aerodynamic degrees of freedom to develop a complete six degree of freedom quasi steady-state aerodynamic model for any vehicle.
Exploring Discretization Error in Simulation-Based Aerodynamic Databases
NASA Technical Reports Server (NTRS)
Aftosmis, Michael J.; Nemec, Marian
2010-01-01
This work examines the level of discretization error in simulation-based aerodynamic databases and introduces strategies for error control. Simulations are performed using a parallel, multi-level Euler solver on embedded-boundary Cartesian meshes. Discretization errors in user-selected outputs are estimated using the method of adjoint-weighted residuals and we use adaptive mesh refinement to reduce these errors to specified tolerances. Using this framework, we examine the behavior of discretization error throughout a token database computed for a NACA 0012 airfoil consisting of 120 cases. We compare the cost and accuracy of two approaches for aerodynamic database generation. In the first approach, mesh adaptation is used to compute all cases in the database to a prescribed level of accuracy. The second approach conducts all simulations using the same computational mesh without adaptation. We quantitatively assess the error landscape and computational costs in both databases. This investigation highlights sensitivities of the database under a variety of conditions. The presence of transonic shocks or the stiffness in the governing equations near the incompressible limit are shown to dramatically increase discretization error requiring additional mesh resolution to control. Results show that such pathologies lead to error levels that vary by over factor of 40 when using a fixed mesh throughout the database. Alternatively, controlling this sensitivity through mesh adaptation leads to mesh sizes which span two orders of magnitude. We propose strategies to minimize simulation cost in sensitive regions and discuss the role of error-estimation in database quality.
Influence of Wake Models on Calculated Tiltrotor Aerodynamics
NASA Technical Reports Server (NTRS)
Johnson, Wayne
2001-01-01
The tiltrotor aircraft configuration has the potential to revolutionize air transportation by providing an economical combination of vertical take-off and landing capability with efficient, high-speed cruise flight. To achieve this potential it is necessary to have validated analytical tools that will support future tiltrotor aircraft development. These analytical tools must calculate tiltrotor aeromechanical behavior, including performance, structural loads, vibration, and aeroelastic stability, with an accuracy established by correlation with measured tiltrotor data. The recent test of the Tilt Rotor Aeroacoustic Model (TRAM) with a single,l/4-scale V-22 rotor in the German-Dutch Wind Tunnel (DNW) provides an extensive set of aeroacoustic, performance, and structural loads data. This paper will examine the influence of wake models on calculated tiltrotor aerodynamics, comparing calculations of performance and airloads with TRAM DNW measurements. The calculations will be performed using the comprehensive analysis CAMRAD II.
NASA Technical Reports Server (NTRS)
Morelli, E. A.; Proffitt, M. S.
1999-01-01
The data for longitudinal non-dimensional, aerodynamic coefficients in the High Speed Research Cycle 2B aerodynamic database were modeled using polynomial expressions identified with an orthogonal function modeling technique. The discrepancy between the tabular aerodynamic data and the polynomial models was tested and shown to be less than 15 percent for drag, lift, and pitching moment coefficients over the entire flight envelope. Most of this discrepancy was traced to smoothing local measurement noise and to the omission of mass case 5 data in the modeling process. A simulation check case showed that the polynomial models provided a compact and accurate representation of the nonlinear aerodynamic dependencies contained in the HSR Cycle 2B tabular aerodynamic database.
NASA Astrophysics Data System (ADS)
Shortis, Mark R.; Robson, Stuart; Jones, Thomas W.; Goad, William K.; Lunsford, Charles B.
2016-06-01
Aerospace engineers require measurements of the shape of aerodynamic surfaces and the six degree of freedom (6DoF) position and orientation of aerospace models to analyse structural dynamics and aerodynamic forces. The measurement technique must be non-contact, accurate, reliable, have a high sample rate and preferably be non-intrusive. Close range photogrammetry based on multiple, synchronised, commercial-off-the-shelf digital cameras can supply surface shape and 6DoF data at 5-15Hz with customisable accuracies. This paper describes data acquisition systems designed and implemented at NASA Langley Research Center to capture surface shapes and 6DoF data. System calibration and data processing techniques are discussed. Examples of experiments and data outputs are described.
Modification of k-ω turbulence model for predicting airfoil aerodynamic performance
NASA Astrophysics Data System (ADS)
Peng, Bo; Yan, Hao; Fang, Hong; Wang, Ming
2015-06-01
Predicting wind turbine S825 airfoil's aerodynamic performance is crucial to improving its energy efficiency and reducing its environmental impact. In this paper, a numerical simulation on the wind turbine S825 airfoil is conducted with k-ω turbulence model at different attack angles. By comparing with experimental data, a new method of modifying k-ω model is proposed. A modifying function is proposed to limit the production term in ω equation based on fluid rotation and deformation. This method improves turbulent viscosity and decreases separating region when the airfoil works at large separating conditions. The predictive accuracy could be improved by using the modified k-ω turbulence model.
Modeling the Launch Abort Vehicle's Subsonic Aerodynamics from Free Flight Testing
NASA Technical Reports Server (NTRS)
Hartman, Christopher L.
2010-01-01
An investigation into the aerodynamics of the Launch Abort Vehicle for NASA's Constellation Crew Launch Vehicle in the subsonic, incompressible flow regime was conducted in the NASA Langley 20-ft Vertical Spin Tunnel. Time histories of center of mass position and Euler Angles are captured using photogrammetry. Time histories of the wind tunnel's airspeed and dynamic pressure are recorded as well. The primary objective of the investigation is to determine models for the aerodynamic yaw and pitch moments that provide insight into the static and dynamic stability of the vehicle. System IDentification Programs for AirCraft (SIDPAC) is used to determine the aerodynamic model structure and estimate model parameters. Aerodynamic models for the aerodynamic body Y and Z force coefficients, and the pitching and yawing moment coefficients were identified.
Modelling Aerodynamically Generated Sound: Recent Advances in Rotor Noise Prediction
NASA Technical Reports Server (NTRS)
Brentner, Kenneth S.
2000-01-01
A great deal of progress has been made in the modeling of aerodynamically generated sound for rotors over the past decade. The Ffowcs Williams-Hawkings (FW-H ) equation has been the foundation for much of the development. Both subsonic and supersonic quadrupole noise formulations have been developed for the prediction of high-speed impulsive noise. In an effort to eliminate the need to compute the quadrupole contribution, the FW-H has also been utilized on permeable surfaces surrounding all physical noise sources. Comparison of the Kirchhoff formulation for moving surfaces with the FW-H equation have shown that the Kirchhoff formulation for moving surfaces can give erroneous results for aeroacoustic problems.
Unsteady Aerodynamic Modeling in Roll for the NASA Generic Transport Model
NASA Technical Reports Server (NTRS)
Murphy, Patrick C.; Klein, Vladislav; Frink, Neal T.
2012-01-01
Reducing the impact of loss-of-control conditions on commercial transport aircraft is a primary goal of the NASA Aviation Safety Program. One aspect in developing the supporting technologies is to improve the aerodynamic models that represent these adverse conditions. Aerodynamic models appropriate for loss of control conditions require a more general mathematical representation to predict nonlinear unsteady behaviors. In this paper, a more general mathematical model is proposed for the subscale NASA Generic Transport Model (GTM) that covers both low and high angles of attack. Particular attention is devoted to the stall region where full-scale transports have demonstrated a tendency for roll instability. The complete aerodynamic model was estimated from dynamic wind-tunnel data. Advanced computational methods are used to improve understanding and visualize the flow physics within the region where roll instability is a factor.
NASA Astrophysics Data System (ADS)
Airoldi, Alessandro; Fournier, Stephane; Borlandelli, Elena; Bettini, Paolo; Sala, Giuseppe
2017-04-01
The paper discusses the approaches for the design and manufacturing of morphing skins based on rectangular-shaped composite corrugated laminates and proposes a novel solution to prevent detrimental effects of corrugation on aerodynamic performances. Additionally, more complex corrugated shapes are presented and analysed. The manufacturing issues related to the production of corrugated laminates are discussed and tests are performed to compare different solutions and to assess the validity of analytical and numerical predictions. The solution presented to develop an aerodynamically efficient skin consists in the integration of an elastomeric cover in the corrugated laminate. The related manufacturing process is presented and assessed, and a fully nonlinear numerical model is developed and characterized to study the behaviour of this skin concept in different load conditions. Finally, configurations based on combinations of individual rectangular-shaped corrugated panels are considered. Their structural properties are numerically investigated by varying geometrical parameters. Performance indices are defined to compare structural stiffness contributions in non-morphing directions with the ones of conventional panels of the same weight. Numerical studies also show that the extension of the concept to complex corrugated shapes may improve both the design flexibility and some specific performances with respect to rectangular shaped corrugations. The overall results validate the design approaches and manufacturing processes to produce corrugated laminates and indicate that the solution for the integration of an elastomeric cover is a feasible and promising method to enhance the aerodynamic efficiency of corrugated skins.
The Benchmark Active Controls Technology Model Aerodynamic Data
NASA Technical Reports Server (NTRS)
Scott, Robert C.; Hoadley, Sherwood T.; Wieseman, Carol D.; Durham, Michael H.
1997-01-01
The Benchmark Active Controls Technology (BACT) model is a part of the Benchmark Models Program (BMP). The BMP is a NASA Langley Research Center program that includes a series of models which were used to study different aeroelastic phenomena and to validate computational fluid dynamics codes. The primary objective of BACT testing was to obtain steady and unsteady loads, accelerations, and aerodynamic pressures due to control surface activity in order to calibrate unsteady CFD codes and active control design tools. Three wind-tunnel tests in the Transonic Dynamics Tunnel (TDT) have been completed. The first and parts of the second and third tests focused on collecting open-loop data to define the model's aeroservoelastic characteristics, including the flutter boundary across the Mach range. It is this data that is being presented in this paper. An extensive database of over 3000 data sets was obtained. This database includes steady and unsteady control surface effectiveness data, including pressure distributions, control surface hinge moments, and overall model loads due to deflections of a trailing edge control surface and upper and lower surface
Validation of aerodynamic parameters at high angles of attack for RAE high incidence research models
NASA Technical Reports Server (NTRS)
Ross, A. Jean; Edwards, Geraldine F.; Klein, Vladislav; Batterson, James G.
1987-01-01
Two series of free-flight tests have been conducted for combat aircraft configuration research models in order to investigate flight behavior near departure conditions as well as to obtain response data from which aerodynamic characteristics can be derived. The structure of the mathematical model and values for the mathematical derivatives have been obtained through an analysis of the first series, using stepwise regression. The results thus obtained are the bases of the design of active control laws. Flight test results for a novel configuration are compared with predicted responses.
NASA Astrophysics Data System (ADS)
Petoshin, V. I.; Chasovnikov, E. A.
2011-05-01
Aerodynamic loads in problems of flight dynamics of passenger aircraft in stalled flow regimes are described using a mathematical model that includes an ordinary linear first-order differential equation. A procedure for determining the parameters of the mathematical model is proposed which is based on approximating experimental frequency characteristics with the frequency characteristics of the linearized mathematical model. The mathematical model was verified by tests of a modern passenger aircraft model in a wind tunnel.
NASA Technical Reports Server (NTRS)
Van Norman, John W.; Dyakonov, Artem; Schoenenberger, Mark; Davis, Jody; Muppidi, Suman; Tang, Chun; Bose, Deepak; Mobley, Brandon; Clark, Ian
2015-01-01
An overview of pre-flight aerodynamic models for the Low Density Supersonic Decelerator (LDSD) Supersonic Flight Dynamics Test (SFDT) campaign is presented, with comparisons to reconstructed flight data and discussion of model updates. The SFDT campaign objective is to test Supersonic Inflatable Aerodynamic Decelerator (SIAD) and large supersonic parachute technologies at high altitude Earth conditions relevant to entry, descent, and landing (EDL) at Mars. Nominal SIAD test conditions are attained by lifting a test vehicle (TV) to 36 km altitude with a large helium balloon, then accelerating the TV to Mach 4 and and 53 km altitude with a solid rocket motor. The first flight test (SFDT-1) delivered a 6 meter diameter robotic mission class decelerator (SIAD-R) to several seconds of flight on June 28, 2014, and was successful in demonstrating the SFDT flight system concept and SIAD-R. The trajectory was off-nominal, however, lofting to over 8 km higher than predicted in flight simulations. Comparisons between reconstructed flight data and aerodynamic models show that SIAD-R aerodynamic performance was in good agreement with pre-flight predictions. Similar comparisons of powered ascent phase aerodynamics show that the pre-flight model overpredicted TV pitch stability, leading to underprediction of trajectory peak altitude. Comparisons between pre-flight aerodynamic models and reconstructed flight data are shown, and changes to aerodynamic models using improved fidelity and knowledge gained from SFDT-1 are discussed.
NASA Technical Reports Server (NTRS)
Alcorn, Charles W.; Britcher, Colin
1988-01-01
An experimental investigation is reported on slanted base ogive cylinders at zero incidence. The Mach number range is 0.05 to 0.3. All flow disturbances associated with wind tunnel supports are eliminated in this investigation by magnetically suspending the wind tunnel models. The sudden and drastic changes in the lift, pitching moment, and drag for a slight change in base slant angle are reported. Flow visualization with liquid crystals and oil is used to observe base flow patterns, which are responsible for the sudden changes in aerodynamic characteristics. Hysteretic effects in base flow pattern changes are present in this investigation and are reported. The effect of a wire support attachment on the 0 deg slanted base model is studied. Computational drag and transition location results using VSAERO and SANDRAG are presented and compared with experimental results. Base pressure measurements over the slanted bases are made with an onboard pressure transducer using remote data telemetry.
Unsteady aerodynamic modeling for arbitrary motions. [for active control techniques
NASA Technical Reports Server (NTRS)
Edwards, J. W.
1977-01-01
Results indicating that unsteady aerodynamic loads derived under the assumption of simple harmonic motions executed by airfoil or wing can be extended to arbitrary motions are summarized. The generalized Theodorsen (1953) function referable to loads due to simple harmonic oscillations of a wing section in incompressible flow, the Laplace inversion integral for unsteady aerodynamic loads, calculations of root loci of aeroelastic loads, and analysis of generalized compressible transient airloads are discussed.
Progressive Aerodynamic Model Identification From Dynamic Water Tunnel Test of the F-16XL Aircraft
NASA Technical Reports Server (NTRS)
Murphy, Patrick C.; Klein, Vladislav; Szyba, Nathan M.
2004-01-01
Development of a general aerodynamic model that is adequate for predicting the forces and moments in the nonlinear and unsteady portions of the flight envelope has not been accomplished to a satisfactory degree. Predicting aerodynamic response during arbitrary motion of an aircraft over the complete flight envelope requires further development of the mathematical model and the associated methods for ground-based testing in order to allow identification of the model. In this study, a general nonlinear unsteady aerodynamic model is presented, followed by a summary of a linear modeling methodology that includes test and identification methods, and then a progressive series of steps suggesting a roadmap to develop a general nonlinear methodology that defines modeling, testing, and identification methods. Initial steps of the general methodology were applied to static and oscillatory test data to identify rolling-moment coefficient. Static measurements uncovered complicated dependencies of the aerodynamic coefficient on angle of attack and sideslip in the stall region making it difficult to find a simple analytical expression for the measurement data. In order to assess the effect of sideslip on the damping and unsteady terms, oscillatory tests in roll were conducted at different values of an initial offset in sideslip. Candidate runs for analyses were selected where higher order harmonics were required for the model and where in-phase and out-of-phase components varied with frequency. From these results it was found that only data in the angle-of-attack range of 35 degrees to 37.5 degrees met these requirements. From the limited results it was observed that the identified models fit the data well and both the damping-in-roll and the unsteady term gain are decreasing with increasing sideslip and motion amplitude. Limited similarity between parameter values in the nonlinear model and the linear model suggest that identifiability of parameters in both terms may be a
NASA Astrophysics Data System (ADS)
Pollock, Michael; Colli, Matteo; Stagnaro, Mattia; Lanza, Luca; Quinn, Paul; Dutton, Mark; O'Donnell, Greg; Wilkinson, Mark; Black, Andrew; O'Connell, Enda
2016-04-01
observed in the vicinity of the collector, compared to the standard gauge shapes. Both the air velocity and the turbulent kinetic energy fields present structures that may improve the interception of particles by the aerodynamic gauge collector. To provide empirical validation, a field-based experimental campaign was undertaken at four UK research stations to compare the results of aerodynamic and conventional gauges, mounted in juxtaposition. The reference measurement is recorded using a rain gauge pit, as specified by the WMO. The results appear to demonstrate how the effect of the wind on rainfall measurements is influenced by the gauge shape and the mounting height. Significant undercatch is observed compared to the reference measurement. Aerodynamic gauges mounted on the ground catch more rainfall than juxtaposed straight-sided gauges, in most instances. This appears to provide some preliminary validation of the CFD model. The indication that an aerodynamic profile improves the gauge catching capability could be confirmed by tracking the hydrometeor trajectories with a Lagrangian method, based on the available set of airflows; and investigating time-dependent aerodynamic features by means of dedicated CFD simulations. Furthermore, wind-tunnel tests could be carried out to provide more robust physical validation of the CFD model.
NASA Astrophysics Data System (ADS)
Xin, Dabo; Ou, Jinping
2007-06-01
Combining the computational fluid dynamics-based numerical simulation with the forced vibration technique for extraction of aerodynamic derivatives, an approach for calculating the aerodynamic derivatives and the critical flutter wind speed for long-span bridges is presented in this paper. The RNG k-ɛ turbulent model is introduced to establish the governing equations, including the continuity equation and the Navier-Stokes equations, for solving the wind flow field around a two-dimensional bridge section. To illustrate the effectiveness and accuracy of the proposed approach, a simple application to the Hume Bridge in China is provided, and the numerical results show that the aerodynamic derivatives and the critical flutter wind speed obtained agree well with the wind tunnel test results.
CAD-Based Aerodynamic Design of Complex Configurations using a Cartesian Method
NASA Technical Reports Server (NTRS)
Nemec, Marian; Aftosmis, Michael J.; Pulliam, Thomas H.
2003-01-01
A modular framework for aerodynamic optimization of complex geometries is developed. By working directly with a parametric CAD system, complex-geometry models are modified nnd tessellated in an automatic fashion. The use of a component-based Cartesian method significantly reduces the demands on the CAD system, and also provides for robust and efficient flowfield analysis. The optimization is controlled using either a genetic or quasi-Newton algorithm. Parallel efficiency of the framework is maintained even when subject to limited CAD resources by dynamically re-allocating the processors of the flow solver. Overall, the resulting framework can explore designs incorporating large shape modifications and changes in topology.
NASA Technical Reports Server (NTRS)
Batterson, J. G.
1986-01-01
The successful parametric modeling of the aerodynamics for an airplane operating at high angles of attack or sideslip is performed in two phases. First the aerodynamic model structure must be determined and second the associated aerodynamic parameters (stability and control derivatives) must be estimated for that model. The purpose of this paper is to document two versions of a stepwise regression computer program which were developed for the determination of airplane aerodynamic model structure and to provide two examples of their use on computer generated data. References are provided for the application of the programs to real flight data. The two computer programs that are the subject of this report, STEP and STEPSPL, are written in FORTRAN IV (ANSI l966) compatible with a CDC FTN4 compiler. Both programs are adaptations of a standard forward stepwise regression algorithm. The purpose of the adaptation is to facilitate the selection of a adequate mathematical model of the aerodynamic force and moment coefficients of an airplane from flight test data. The major difference between STEP and STEPSPL is in the basis for the model. The basis for the model in STEP is the standard polynomial Taylor's series expansion of the aerodynamic function about some steady-state trim condition. Program STEPSPL utilizes a set of spline basis functions.
NASA Technical Reports Server (NTRS)
Prasanth, Ravi K.; Klein, Vladislav; Murphy, Patrick C.; Mehra, Raman K.
2005-01-01
This paper describes model structures and parameter estimation algorithms suitable for the identification of unsteady aerodynamic models from input-output data. The model structures presented are state space models and include linear time-invariant (LTI) models and linear parameter-varying (LPV) models. They cover a wide range of local and parameter dependent identification problems arising in unsteady aerodynamics and nonlinear flight dynamics. We present a residue algorithm for estimating model parameters from data. The algorithm can incorporate apriori information and is described in detail. The algorithms are evaluated on the F-16XL wind-tunnel test data from NAS Langley Research Center. Results of numerical evaluation are presented. The paper concludes with a discussion major issues and directions for future work.
NASA Astrophysics Data System (ADS)
Azcona, José; Bouchotrouch, Faisal; González, Marta; Garciandía, Joseba; Munduate, Xabier; Kelberlau, Felix; Nygaard, Tor A.
2014-06-01
Wave tank testing of scaled models is standard practice during the development of floating wind turbine platforms for the validation of the dynamics of conceptual designs. Reliable recreation of the dynamics of a full scale floating wind turbine by a scaled model in a basin requires the precise scaling of the masses and inertias and also the relevant forces and its frequencies acting on the system. The scaling of floating wind turbines based on the Froude number is customary for basin experiments. This method preserves the hydrodynamic similitude, but the resulting Reynolds number is much lower than in full scale. The aerodynamic loads on the rotor are therefore out of scale. Several approaches have been taken to deal with this issue, like using a tuned drag disk or redesigning the scaled rotor. This paper describes the implementation of an alternative method based on the use of a ducted fan located at the model tower top in the place of the rotor. The fan can introduce a variable force that represents the total wind thrust by the rotor. A system controls this force by varying the rpm, and a computer simulation of the full scale rotor provides the desired thrust to be introduced by the fan. This simulation considers the wind turbine control, gusts, turbulent wind, etc. The simulation is performed in synchronicity with the test and it is fed in real time by the displacements and velocities of the platform captured by the acquisition system. Thus, the simulation considers the displacements of the rotor within the wind field and the calculated thrust models the effect of the aerodynamic damping. The system is not able currently to match the effect of gyroscopic momentum. The method has been applied during a test campaign of a semisubmersible platform with full catenary mooring lines for a 6MW wind turbine in scale 1/40 at Ecole Centrale de Nantes. Several tests including pitch free decay under constant wind and combined wave and wind cases have been performed. Data
NASA Technical Reports Server (NTRS)
Truscott, Starr; Parkinson, J B; Ebert, John W , Jr; Valentine, E Floyd
1938-01-01
The present tests illustrate how the aerodynamic drag of a flying boat hull may be reduced by following closely the form of a low drag aerodynamic body and the manner in which the extent of the aerodynamic refinement is limited by poorer hydrodynamic performance. This limit is not sharply defined but is first evidenced by an abnormal flow of water over certain parts of the form accompanied by a sharp increase in resistance. In the case of models 74-A and 75, the resistance (sticking) occurs only at certain combinations of speed, load, and trim and can be avoided by proper control of the trim at high water speeds. Model 75 has higher water resistance at very high speeds than does model 74-A. With constant speed propellers and high takeoff speeds, it appears that the form of model 75 would give slightly better takeoff performance. Model 74-A, however, has lower aerodynamic drag than does model 75 for the same volume of hull.
Gradient-Based Aerodynamic Shape Optimization Using ADI Method for Large-Scale Problems
NASA Technical Reports Server (NTRS)
Pandya, Mohagna J.; Baysal, Oktay
1997-01-01
A gradient-based shape optimization methodology, that is intended for practical three-dimensional aerodynamic applications, has been developed. It is based on the quasi-analytical sensitivities. The flow analysis is rendered by a fully implicit, finite volume formulation of the Euler equations.The aerodynamic sensitivity equation is solved using the alternating-direction-implicit (ADI) algorithm for memory efficiency. A flexible wing geometry model, that is based on surface parameterization and platform schedules, is utilized. The present methodology and its components have been tested via several comparisons. Initially, the flow analysis for for a wing is compared with those obtained using an unfactored, preconditioned conjugate gradient approach (PCG), and an extensively validated CFD code. Then, the sensitivities computed with the present method have been compared with those obtained using the finite-difference and the PCG approaches. Effects of grid refinement and convergence tolerance on the analysis and shape optimization have been explored. Finally the new procedure has been demonstrated in the design of a cranked arrow wing at Mach 2.4. Despite the expected increase in the computational time, the results indicate that shape optimization, which require large numbers of grid points can be resolved with a gradient-based approach.
Mathematical modeling of the aerodynamics of high-angle-of-attack maneuvers
NASA Technical Reports Server (NTRS)
Schiff, L. B.; Tobak, M.; Malcolm, G. N.
1980-01-01
This paper is a review of the current state of aerodynamic mathematical modeling for aircraft motions at high angles of attack. The mathematical model serves to define a set of characteristic motions from whose known aerodynamic responses the aerodynamic response to an arbitrary high angle-of-attack flight maneuver can be predicted. Means are explored of obtaining stability parameter information in terms of the characteristic motions, whether by wind-tunnel experiments, computational methods, or by parameter-identification methods applied to flight-test data. A rationale is presented for selecting and verifying the aerodynamic mathematical model at the lowest necessary level of complexity. Experimental results describing the wing-rock phenomenon are shown to be accommodated within the most recent mathematical model by admitting the existence of aerodynamic hysteresis in the steady-state variation of the rolling moment with roll angle. Interpretation of the experimental results in terms of bifurcation theory reveals the general conditions under which aerodynamic hysteresis must exist.
NASA Astrophysics Data System (ADS)
Liu, Z. Y. C.; Shirzaei, M.
2015-12-01
Impact craters on the terrestrial planets are typically surrounded by a continuous ejecta blanket that the initial emplacement is via ballistic sedimentation. Following an impact event, a significant volume of material is ejected and falling debris surrounds the crater. Aerodynamics rule governs the flight path and determines the spatial distribution of these ejecta. Thus, for the planets with atmosphere, the preserved ejecta deposit directly recorded the interaction of ejecta and atmosphere at the time of impact. In this study, we develop a new framework to establish links between distribution of the ejecta, age of the impact and the properties of local atmosphere. Given the radial distance of the continuous ejecta extent from crater, an inverse aerodynamic modeling approach is employed to estimate the local atmospheric drags and density as well as the lift forces at the time of impact. Based on earlier studies, we incorporate reasonable value ranges for ejection angle, initial velocity, aerodynamic drag, and lift in the model. In order to solve the trajectory differential equations, obtain the best estimate of atmospheric density, and the associated uncertainties, genetic algorithm is applied. The method is validated using synthetic data sets as well as detailed maps of impact ejecta associated with five fresh martian and two lunar impact craters, with diameter of 20-50 m, 10-20 m, respectively. The estimated air density for martian carters range 0.014-0.028 kg/m3, consistent with the recent surface atmospheric density measurement of 0.015-0.020 kg/m3. This constancy indicates the robustness of the presented methodology. In the following, the inversion results for the lunar craters yield air density of 0.003-0.008 kg/m3, which suggest the inversion results are accurate to the second decimal place. This framework will be applied to older martian craters with preserved ejecta blankets, which expect to constrain the long-term evolution of martian atmosphere.
Computations of Aerodynamic Performance Databases Using Output-Based Refinement
NASA Technical Reports Server (NTRS)
Nemec, Marian; Aftosmis, Michael J.
2009-01-01
Objectives: Handle complex geometry problems; Control discretization errors via solution-adaptive mesh refinement; Focus on aerodynamic databases of parametric and optimization studies: 1. Accuracy: satisfy prescribed error bounds 2. Robustness and speed: may require over 105 mesh generations 3. Automation: avoid user supervision Obtain "expert meshes" independent of user skill; and Run every case adaptively in production settings.
Modeling aerodynamic discontinuities and the onset of chaos in flight dynamical systems
NASA Technical Reports Server (NTRS)
Tobak, M.; Chapman, G. T.; Uenal, A.
1986-01-01
Various representations of the aerodynamic contribution to the aircraft's equation of motion are shown to be compatible within the common assumption of their Frechet differentiability. Three forms of invalidating Frechet differentiality are identified, and the mathematical model is amended to accommodate their occurrence. Some of the ways in which chaotic behavior may emerge are discussed, first at the level of the aerodynamic contribution to the equation of motion, and then at the level of the equations of motion themselves.
A simple analytical aerodynamic model of Langley Winged-Cone Aerospace Plane concept
NASA Technical Reports Server (NTRS)
Pamadi, Bandu N.
1994-01-01
A simple three DOF analytical aerodynamic model of the Langley Winged-Coned Aerospace Plane concept is presented in a form suitable for simulation, trajectory optimization, and guidance and control studies. The analytical model is especially suitable for methods based on variational calculus. Analytical expressions are presented for lift, drag, and pitching moment coefficients from subsonic to hypersonic Mach numbers and angles of attack up to +/- 20 deg. This analytical model has break points at Mach numbers of 1.0, 1.4, 4.0, and 6.0. Across these Mach number break points, the lift, drag, and pitching moment coefficients are made continuous but their derivatives are not. There are no break points in angle of attack. The effect of control surface deflection is not considered. The present analytical model compares well with the APAS calculations and wind tunnel test data for most angles of attack and Mach numbers.
The DaveMLTranslator: An Interface for DAVE-ML Aerodynamic Models
NASA Technical Reports Server (NTRS)
Hill, Melissa A.; Jackson, E. Bruce
2007-01-01
It can take weeks or months to incorporate a new aerodynamic model into a vehicle simulation and validate the performance of the model. The Dynamic Aerospace Vehicle Exchange Markup Language (DAVE-ML) has been proposed as a means to reduce the time required to accomplish this task by defining a standard format for typical components of a flight dynamic model. The purpose of this paper is to describe an object-oriented C++ implementation of a class that interfaces a vehicle subsystem model specified in DAVE-ML and a vehicle simulation. Using the DaveMLTranslator class, aerodynamic or other subsystem models can be automatically imported and verified at run-time, significantly reducing the elapsed time between receipt of a DAVE-ML model and its integration into a simulation environment. The translator performs variable initializations, data table lookups, and mathematical calculations for the aerodynamic build-up, and executes any embedded static check-cases for verification. The implementation is efficient, enabling real-time execution. Simple interface code for the model inputs and outputs is the only requirement to integrate the DaveMLTranslator as a vehicle aerodynamic model. The translator makes use of existing table-lookup utilities from the Langley Standard Real-Time Simulation in C++ (LaSRS++). The design and operation of the translator class is described and comparisons with existing, conventional, C++ aerodynamic models of the same vehicle are given.
Calculation of the Aerodynamic Behavior of the Tilt Rotor Aeroacoustic Model (TRAM) in the DNW
NASA Technical Reports Server (NTRS)
Johnson, Wayne
2001-01-01
Comparisons of measured and calculated aerodynamic behavior of a tiltrotor model are presented. The test of the Tilt Rotor Aeroacoustic Model (TRAM) with a single, 1/4-scale V- 22 rotor in the German-Dutch Wind Tunnel (DNW) provides an extensive set of aeroacoustic, performance, and structural loads data. The calculations were performed using the rotorcraft comprehensive analysis CAMRAD II. Presented are comparisons of measured and calculated performance and airloads for helicopter mode operation, as well as calculated induced and profile power. An aerodynamic and wake model and calculation procedure that reflects the unique geometry and phenomena of tiltrotors has been developed. There are major differences between this model and the corresponding aerodynamic and wake model that has been established for helicopter rotors. In general, good correlation between measured and calculated performance and airloads behavior has been shown. Two aspects of the analysis that clearly need improvement are the stall delay model and the trailed vortex formation model.
NASA Technical Reports Server (NTRS)
Goldstein, M.; Rosenbaum, B.
1973-01-01
A model based on Lighthill's theory for predicting aerodynamic noise from a turbulent shear flow is developed. This model is a generalization of the one developed by Ribner. It does not require that the turbulent correlations factor into space and time-dependent parts. It replaces his assumption of isotropic turbulence by the more realistic one of axisymmetric turbulence. In the course of the analysis, a hierarchy of equations is developed wherein each succeeding equation involves more assumptions than the preceding equation but requires less experimental information for its use. The implications of the model for jet noise are discussed. It is shown that for the particular turbulence data considered anisotropy causes the high-frequency self-noise to be beamed downstream.
Application of CFD techniques toward the validation of nonlinear aerodynamic models
NASA Technical Reports Server (NTRS)
Schiff, L. B.; Katz, J.
1985-01-01
Applications of Computational fluid dynamics (CFD) methods to determine the regimes of applicability of nonlinear models describing the unsteady aerodynamic responses to aircraft flight motions are described. The potential advantages of computational methods over experimental methods are discussed and the concepts underlying mathematical modeling are reviewed. The economic and conceptual advantages of the modeling procedure over coupled, simultaneous solutions of the gasdynamic equations and the vehicle's kinematic equations of motion are discussed. The modeling approach, when valid, eliminates the need for costly repetitive computation of flow field solutions. For the test cases considered, the aerodynamic modeling approach is shown to be valid.
NASA Technical Reports Server (NTRS)
Nelson, D. P.; Morris, P. M.
1980-01-01
The component detail design drawings of the one sixth scale model of the variable cycle engine testbed demonstrator exhaust syatem tested are presented. Also provided are the basic acoustic and aerodynamic data acquired during the experimental model tests. The model drawings, an index to the acoustic data, an index to the aerodynamic data, tabulated and graphical acoustic data, and the tabulated aerodynamic data and graphs are discussed.
Nonlinear Aerodynamic Modeling From Flight Data Using Advanced Piloted Maneuvers and Fuzzy Logic
NASA Technical Reports Server (NTRS)
Brandon, Jay M.; Morelli, Eugene A.
2012-01-01
Results of the Aeronautics Research Mission Directorate Seedling Project Phase I research project entitled "Nonlinear Aerodynamics Modeling using Fuzzy Logic" are presented. Efficient and rapid flight test capabilities were developed for estimating highly nonlinear models of airplane aerodynamics over a large flight envelope. Results showed that the flight maneuvers developed, used in conjunction with the fuzzy-logic system identification algorithms, produced very good model fits of the data, with no model structure inputs required, for flight conditions ranging from cruise to departure and spin conditions.
Real-Time Global Nonlinear Aerodynamic Modeling for Learn-To-Fly
NASA Technical Reports Server (NTRS)
Morelli, Eugene A.
2016-01-01
Flight testing and modeling techniques were developed to accurately identify global nonlinear aerodynamic models for aircraft in real time. The techniques were developed and demonstrated during flight testing of a remotely-piloted subscale propeller-driven fixed-wing aircraft using flight test maneuvers designed to simulate a Learn-To-Fly scenario. Prediction testing was used to evaluate the quality of the global models identified in real time. The real-time global nonlinear aerodynamic modeling algorithm will be integrated and further tested with learning adaptive control and guidance for NASA Learn-To-Fly concept flight demonstrations.
NASA Technical Reports Server (NTRS)
Springer, Anthony M.
1998-01-01
The aerodynamic characteristics of two similar models of a lifting body configuration were run in two transonic wind tunnels, one a 16 foot the other a 14-inch and are compared. The 16 foot test used a 2% model while the 14-inch test used a 0.7% scale model. The wind tunnel model configurations varied only in vertical tail size and an aft sting shroud. The results from these two tests compare the effect of tunnel size, Reynolds number, dynamic pressure and blockage on the longitudinal aerodynamic characteristics of the vehicle. The data accuracy and uncertainty are also presented. It was concluded from these tests that the data resultant from a small wind tunnel compares very well to that of a much larger wind tunnel in relation to total vehicle aerodynamic characteristics.
CFD modelling of the aerodynamic effect of trees on urban air pollution dispersion.
Amorim, J H; Rodrigues, V; Tavares, R; Valente, J; Borrego, C
2013-09-01
The current work evaluates the impact of urban trees over the dispersion of carbon monoxide (CO) emitted by road traffic, due to the induced modification of the wind flow characteristics. With this purpose, the standard flow equations with a kε closure for turbulence were extended with the capability to account for the aerodynamic effect of trees over the wind field. Two CFD models were used for testing this numerical approach. Air quality simulations were conducted for two periods of 31h in selected areas of Lisbon and Aveiro, in Portugal, for distinct relative wind directions: approximately 45° and nearly parallel to the main avenue, respectively. The statistical evaluation of modelling performance and uncertainty revealed a significant improvement of results with trees, as shown by the reduction of the NMSE from 0.14 to 0.10 in Lisbon, and from 0.14 to 0.04 in Aveiro, which is independent from the CFD model applied. The consideration of the plant canopy allowed to fulfil the data quality objectives for ambient air quality modelling established by the Directive 2008/50/EC, with an important decrease of the maximum deviation between site measurements and CFD results. In the non-aligned wind situation an average 12% increase of the CO concentrations in the domain was observed as a response to the aerodynamic action of trees over the vertical exchange rates of polluted air with the above roof-level atmosphere; while for the aligned configuration an average 16% decrease was registered due to the enhanced ventilation of the street canyon. These results show that urban air quality can be optimised based on knowledge-based planning of green spaces.
Enhanced ground-based vibration testing for aerodynamic environments
NASA Astrophysics Data System (ADS)
Daborn, P. M.; Ind, P. R.; Ewins, D. J.
2014-12-01
Typical methods of replicating aerodynamic environments in the laboratory are generally poor. A structure which flies "freely" in its normal operating environment, excited over its entire external surface by aerodynamic forces and in all directions simultaneously, is then subjected to a vibration test in the laboratory whilst rigidly attached to a high impedance shaker and excited by forces applied through a few attachment points and in one direction only. The two environments could hardly be more different. The majority of vibration testing is carried out at commercial establishments and it is understandable that little has been published which demonstrates the limitations with the status quo. The primary objective of this research is to do just that with a view to identifying significant improvements in vibration testing in light of modern technology. In this paper, case studies are presented which highlight some of the limitations with typical vibration tests showing that they can lead to significant overtests, sometimes by many orders of magnitude, with the level of overtest varying considerably across a wide range of frequencies. This research shows that substantial benefits can be gained by "freely" suspending the structure in the laboratory and exciting it with a relatively small number of electrodynamic shakers using Multi-Input-Multi-Output (MIMO) control technology. The shaker configuration can be designed to excite the modes within the bandwidth utilising the inherent amplification of the resonances to achieve the desired response levels. This free-free MIMO vibration test approach is shown to result in substantial benefits that include extremely good replication of the aerodynamic environment and significant savings in time as all axes are excited simultaneously instead of the sequential X, Y and Z testing required with traditional vibration tests. In addition, substantial cost savings can be achieved by replacing some expensive large shaker systems
A comprehensive analytical model of rotorcraft aerodynamics and dynamics. Part 3: Program manual
NASA Technical Reports Server (NTRS)
Johnson, W.
1980-01-01
The computer program for a comprehensive analytical model of rotorcraft aerodynamics and dynamics is described. This analysis is designed to calculate rotor performance, loads, and noise; the helicopter vibration and gust response; the flight dynamics and handling qualities; and the system aeroelastic stability. The analysis is a combination of structural, inertial, and aerodynamic models that is applicable to a wide range of problems and a wide class of vehicles. The analysis is intended for use in the design, testing, and evaluation of rotors and rotorcraft and to be a basis for further development of rotary wing theories.
Aerodynamic and acoustic test of a United Technologies model scale rotor at DNW
NASA Technical Reports Server (NTRS)
Yu, Yung H.; Liu, Sandy R.; Jordan, Dave E.; Landgrebe, Anton J.; Lorber, Peter F.; Pollack, Michael J.; Martin, Ruth M.
1990-01-01
The UTC model scale rotors, the DNW wind tunnel, the AFDD rotary wing test stand, the UTRC and AFDD aerodynamic and acoustic data acquisition systems, and the scope of test matrices are discussed and an introduction to the test results is provided. It is pointed out that a comprehensive aero/acoustic database of several configurations of the UTC scaled model rotor has been created. The data is expected to improve understanding of rotor aerodynamics, acoustics, and dynamics, and lead to enhanced analytical methodology and design capabilities for the next generation of rotorcraft.
Investigation of Tractor Base Bleeding for Heavy Vehicle Aerodynamic Drag Reduction
Ortega, J; Salari, K; Storms, B
2007-10-25
One of the main contributors to the aerodynamic drag of a heavy vehicle is tractor-trailer gap drag, which arises when the vehicle operates within a crosswind. Under this operating condition, freestream flow is entrained into the tractor-trailer gap, imparting a momentum exchange to the vehicle and subsequently increasing the aerodynamic drag. While a number of add-on devices, including side extenders, splitter plates, vortex stabilizers, and gap sealers, have been previously tested to alleviate this source of drag, side extenders remain the primary add-on device of choice for reducing tractor-trailer gap drag. However, side extenders are not without maintenance and operational issues. When a heavy vehicle pivots sharply with respect to the trailer, as can occur during loading or unloading operations, the side extenders can become crushed against the trailer. Consequently, fleet operators are forced to incur additional costs to cover the repair or replacement of the damaged side extenders. This issue can be overcome by either shortening the side extenders or by devising an alternative drag reduction concept that can perform just as effectively as side extenders. To explore such a concept, we investigate tractor base bleeding as a means of reducing gap drag. Wind tunnel measurements are made on a 1:20 scale heavy vehicle model at a vehicle width-based Reynolds number of 420,000. The tractor bleeding flow, which is delivered through a porous material embedded within the tractor base, is introduced into the tractor-trailer gap at bleeding coefficients ranging from 0.0-0.018. To determine the performance of tractor base bleeding under more realistic operating conditions, computational fluid dynamics simulations are performed on a full-scale heavy vehicle within a crosswind for bleeding coefficients ranging from 0.0-0.13.
Post-Stall Aerodynamic Modeling and Gain-Scheduled Control Design
NASA Technical Reports Server (NTRS)
Wu, Fen; Gopalarathnam, Ashok; Kim, Sungwan
2005-01-01
A multidisciplinary research e.ort that combines aerodynamic modeling and gain-scheduled control design for aircraft flight at post-stall conditions is described. The aerodynamic modeling uses a decambering approach for rapid prediction of post-stall aerodynamic characteristics of multiple-wing con.gurations using known section data. The approach is successful in bringing to light multiple solutions at post-stall angles of attack right during the iteration process. The predictions agree fairly well with experimental results from wind tunnel tests. The control research was focused on actuator saturation and .ight transition between low and high angles of attack regions for near- and post-stall aircraft using advanced LPV control techniques. The new control approaches maintain adequate control capability to handle high angle of attack aircraft control with stability and performance guarantee.
Energy-Based Design Methodology for Air Vehicle Systems: Aerodynamic Correlation Study
2005-03-01
ENERGY -BASED DESIGN METHODOLOGY FOR AIR VEHICLE SYSTEMS : AERODYNAMIC CORRELATION STUDY AFOSR: FA9550-64-"t/Dr. John Schmisseur AFOSR-NA C>(4-1-0- I...drag estimation and vehicle-level utilization of energy . The exergy utilization of a wing in a steady, low subsonic, three-dimensional, viscous flow...5a. CONTRACT NUMBER Energy -Based Design Methodology For Air Vehicle 5b. GRANT NUMBER Systems : Aerodynamic Correlation Study FA9550,-64 (9 4-1-- !(1 5c
Helicopter flight dynamics simulation with refined aerodynamic modeling
NASA Astrophysics Data System (ADS)
Theodore, Colin Rhys
This dissertation describes the development of a coupled rotor-fuselage flight dynamic simulation that includes a maneuvering free wake model and a coupled flap-lag-torsion flexible blade representation. This mathematical model is used to investigate effects of main rotor inflow and blade modeling on various flight dynamics characteristics for both articulated and hingeless rotor helicopters. The inclusion of the free wake model requires the development of new numerical procedures for the calculation of trim equilibrium positions, for the extraction of high-order, constant coefficient linearized models, and for the calculation of the free flight responses to arbitrary pilot inputs. The free wake model, previously developed by other investigators at the University of Maryland, is capable of modeling the changes in rotor wake geometry resulting from maneuvers, and the effects of such changes on the main rotor inflow. The overall flight dynamic model is capable of simulating the helicopter behavior during maneuvers that can be arbitrarily large. The combination of sophisticated models of rotor wake and blade flexibility enables the flight dynamics model to capture the effects of maneuvers with unprecedented accuracy for simulations based on first principles: this is the main contribution of the research presented in this dissertation. The increased accuracy brought about by the free wake model significantly improves the predictions of the helicopter trim state for both helicopter configurations considered in this study. This is especially true in low speed flight and hover. The most significant improvements are seen in the predictions of the main rotor collective and power required by the rotor, which can be significantly underpredicted using traditional linear inflow models. Results show that the free-flight on-axis responses to pilot inputs can be predicted with good accuracy with a relatively unsophisticated models that do not include either a free wake nor a
NASA Technical Reports Server (NTRS)
Petot, D.; Loiseau, H.
1982-01-01
Unsteady aerodynamic methods adopted for the study of aeroelasticity in helicopters are considered with focus on the development of a semiempirical model of unsteady aerodynamic forces acting on an oscillating profile at high incidence. The successive smoothing algorithm described leads to the model's coefficients in a very satisfactory manner.
A faster optimization method based on support vector regression for aerodynamic problems
NASA Astrophysics Data System (ADS)
Yang, Xixiang; Zhang, Weihua
2013-09-01
In this paper, a new strategy for optimal design of complex aerodynamic configuration with a reasonable low computational effort is proposed. In order to solve the formulated aerodynamic optimization problem with heavy computation complexity, two steps are taken: (1) a sequential approximation method based on support vector regression (SVR) and hybrid cross validation strategy, is proposed to predict aerodynamic coefficients, and thus approximates the objective function and constraint conditions of the originally formulated optimization problem with given limited sample points; (2) a sequential optimization algorithm is proposed to ensure the obtained optimal solution by solving the approximation optimization problem in step (1) is very close to the optimal solution of the originally formulated optimization problem. In the end, we adopt a complex aerodynamic design problem, that is optimal aerodynamic design of a flight vehicle with grid fins, to demonstrate our proposed optimization methods, and numerical results show that better results can be obtained with a significantly lower computational effort than using classical optimization techniques.
NASA Astrophysics Data System (ADS)
Cheng, Bo; Hu, Zheng; Deng, Xinyan
2010-11-01
Body movements of flying animals change their effective wing kinematics and influence aerodynamic forces. Our previous studies found that substantial aerodynamic damping was produced by flapping wings during body rotation through a passive mechanism we termed flapping counter-torque (FCT). Here we present the aerodynamic damping produced by flapping wings during body translations, which we termed flapping counter-forces (FCFs). Analytical models were derived and the aerodynamic effect of spanwise flow and wing-wake interaction were also explored. The FCFs are dependent on body velocities, wing beat amplitude and frequency. Aerodynamic force and PIV measurements were compared with the analytical models. The experiments were conducted on a pair of dynamically scaled robotic model wings in an oil tank. Experiments in air using a pair of high frequency flapping wing further validate the models. Complete 6-DOF flight dynamic model was derived.
Structural Verification and Modeling of a Tension Cone Inflatable Aerodynamic Decelerator
NASA Technical Reports Server (NTRS)
Tanner, Christopher L.; Cruz, Juan R.; Braun, Robert D.
2010-01-01
Verification analyses were conducted on membrane structures pertaining to a tension cone inflatable aerodynamic decelerator using the analysis code LS-DYNA. The responses of three structures - a cylinder, torus, and tension shell - were compared against linear theory for various loading cases. Stress distribution, buckling behavior, and wrinkling behavior were investigated. In general, agreement between theory and LS-DYNA was very good for all cases investigated. These verification cases exposed the important effects of using a linear elastic liner in membrane structures under compression. Finally, a tension cone wind tunnel test article is modeled in LS-DYNA for which preliminary results are presented. Unlike data from supersonic wind tunnel testing, the segmented tension shell and torus experienced oscillatory behavior when subjected to a steady aerodynamic pressure distribution. This work is presented as a work in progress towards development of a fluid-structures interaction mechanism to investigate aeroelastic behavior of inflatable aerodynamic decelerators.
NASA Technical Reports Server (NTRS)
Mann, M. J.; Langhans, R. A.
1977-01-01
The effects of wing trailing-edge control surfaces on the static transonic aerodynamic characteristics of a transport configuration with a supercritical wing were studied. The configuration was tested with both an area-ruled fuselage and a cylindrical fuselage. The Mach number range was from 0.80 to 0.96 and the angle of attack range was from -1 deg to 12 deg. The Reynolds number was 1,580,000 based on the mean aerodynamic chord. Tabular data are presented.
NASA Astrophysics Data System (ADS)
Mohrfeld-Halterman, J. A.; Uddin, M.
2016-07-01
We described in this paper the development of a high fidelity vehicle aerodynamic model to fit wind tunnel test data over a wide range of vehicle orientations. We also present a comparison between the effects of this proposed model and a conventional quasi steady-state aerodynamic model on race vehicle simulation results. This is done by implementing both of these models independently in multi-body quasi steady-state simulations to determine the effects of the high fidelity aerodynamic model on race vehicle performance metrics. The quasi steady state vehicle simulation is developed with a multi-body NASCAR Truck vehicle model, and simulations are conducted for three different types of NASCAR race tracks, a short track, a one and a half mile intermediate track, and a higher speed, two mile intermediate race track. For each track simulation, the effects of the aerodynamic model on handling, maximum corner speed, and drive force metrics are analysed. The accuracy of the high-fidelity model is shown to reduce the aerodynamic model error relative to the conventional aerodynamic model, and the increased accuracy of the high fidelity aerodynamic model is found to have realisable effects on the performance metric predictions on the intermediate tracks resulting from the quasi steady-state simulation.
Salari, K; Ortega, J
2010-12-13
Lawrence Livermore National Laboratory (LLNL) as part of its Department of Energy (DOE), Energy Efficiency and Renewable Energy (EERE), and Vehicle Technologies Program (VTP) effort has investigated class 8 tractor-trailer aerodynamics for many years. This effort has identified many drag producing flow structures around the heavy vehicles and also has designed and tested many new active and passive drag reduction techniques and concepts for significant on the road fuel economy improvements. As part of this effort a database of experimental, computational, and conceptual design for aerodynamic drag reduction devices has been established. The objective of this report is to provide design guidance for trailer base devices to improve their aerodynamic performance. These devices are commonly referred to as boattails, base flaps, tail devices, and etc. The information provided here is based on past research and our most recent full-scale experimental investigations in collaboration with Navistar Inc. Additional supporting data from LLNL/Navistar wind tunnel, track test, and on the road test will be published soon. The trailer base devices can be identified by 4 flat panels that are attached to the rear edges of the trailer base to form a closed cavity. These devices have been engineered in many different forms such as, inflatable and non-inflatable, 3 and 4-sided, closed and open cavity, and etc. The following is an in-depth discussion with some recommendations, based on existing data and current research activities, of changes that could be made to these devices to improve their aerodynamic performance. There are 6 primary factors that could influence the aerodynamic performance of trailer base devices: (1) Deflection angle; (2) Boattail length; (3) Sealing of edges and corners; (4) 3 versus 4-sided, Position of the 4th plate; (5) Boattail vertical extension, Skirt - boattail transition; and (6) Closed versus open cavity.
Towards a predictive vortex model for 2D non-linear aerodynamics
NASA Astrophysics Data System (ADS)
Darakananda, Darwin; Eldredge, Jeff D.
2014-11-01
In previous work (Hemati et al 2014), we presented a framework in which a low-order point vortex model can be optimized to capture the non-linear aerodynamics of a wing undergoing arbitrary rigid body motion. Rather than determine the time-varying vortex strengths with the Kutta condition, these strengths were chosen to minimize the difference between the force predicted by the model and pre-existing empirical data. Here, we present ongoing extensions of this model. With the help of tools from dynamical systems theory, we develop a means to incrementally optimize the model against new data. This opens the possibility for using the model in a dynamic estimator context. Self-sustained vortex shedding from wings is achieved using a criterion based on the leading edge suction parameter. We demonstrate the model on a variety of canonical problems, including pitch-up, oscillatory heaving and pitching, and impulsive translation of a plate at various angles of attack. This work has been supported by AFOSR, under Award FA9550-11-1-0098.
Computational modeling of aerodynamics in the fast forward flight of hummingbirds
NASA Astrophysics Data System (ADS)
Song, Jialei; Luo, Haoxiang; Tobalske, Bret; Hedrick, Tyson
2015-11-01
Computational models of the hummingbird at flight speed 8.3 m/s is built based on high-speed imaging of the real bird flight in the wind tunnel. The goal is to understand the lift and thrust production of the wings at the high advance ratio (flight speed to the average wingtip speed) around 1. Both the full 3D CFD model based on an immersed-boundary method and the blade-element model based on quasi-steady flow assumption were adopted to analyze the aerodynamics. The result shows that while the weight support is generated during downstroke, little negative weight support is produced during upstroke. On the other hand, thrust is generated during both downstroke and upstroke, which allows the bird to overcome drag induced at fast flight. The lift and thrust characteristics are closely related to the instantaneous wing position and motion. In addition, the flow visualization shows that the leading-edge vortex is stable during most of the wing-beat, which may have contributed to the lift and thrust enhancement. NSF CBET-0954381.
NASA Astrophysics Data System (ADS)
Menicovich, David
material and energy consumption profiles of tall building. To date, the increasing use of light-weight and high-strength materials in tall buildings, with greater flexibility and reduced damping, has increased susceptibility to dynamic wind load effects that limit the gains afforded by incorporating these new materials. Wind, particularly fluctuating wind and its interaction with buildings induces two main responses; alongwind - in the direction of the flow and crosswind - perpendicular to the flow. The main risk associated with this vulnerability is resonant oscillations induced by von-Karman-like vortex shedding at or near the natural frequency of the structure caused by flow separation. Dynamic wind loading effects often increase with a power of wind speed greater than 3, thus increasingly, tall buildings pay a significant price in material to increase the natural frequency and/or the damping to overcome this response. In particular, crosswind response often governs serviceability (human habitability) design criteria of slender buildings. Currently, reducing crosswind response relies on a Solid-based Aerodynamic Modification (SAM), either by changing structural or geometric characteristics such as the tower shape or through the addition of damping systems. While this approach has merit it has two major drawbacks: firstly, the loss of valuable rentable areas and high construction costs due to increased structural requirements for mass and stiffness, further contributing towards the high consumption of non-renewable resources by the commercial building sector. For example, in order to insure human comfort within an acceptable range of crosswind response induced accelerations at the top of a building, an aerodynamically efficient plan shape comes at the expense of floor area. To compensate for the loss of valuable area compensatory stories are required, resulting in an increase in wind loads and construction costs. Secondly, a limited, if at all, ability to adaptively
Technology Transfer Automated Retrieval System (TEKTRAN)
Application of the Two-Source Energy Balance (TSEB) Model using land surface temperature (LST) requires aerodynamic resistance parameterizations for the flux exchange above the canopy layer, within the canopy air space and at the soil/substrate surface. There are a number of aerodynamic resistance f...
Optimal cycling time trial position models: aerodynamics versus power output and metabolic energy.
Fintelman, D M; Sterling, M; Hemida, H; Li, F-X
2014-06-03
The aerodynamic drag of a cyclist in time trial (TT) position is strongly influenced by the torso angle. While decreasing the torso angle reduces the drag, it limits the physiological functioning of the cyclist. Therefore the aims of this study were to predict the optimal TT cycling position as function of the cycling speed and to determine at which speed the aerodynamic power losses start to dominate. Two models were developed to determine the optimal torso angle: a 'Metabolic Energy Model' and a 'Power Output Model'. The Metabolic Energy Model minimised the required cycling energy expenditure, while the Power Output Model maximised the cyclists׳ power output. The input parameters were experimentally collected from 19 TT cyclists at different torso angle positions (0-24°). The results showed that for both models, the optimal torso angle depends strongly on the cycling speed, with decreasing torso angles at increasing speeds. The aerodynamic losses outweigh the power losses at cycling speeds above 46km/h. However, a fully horizontal torso is not optimal. For speeds below 30km/h, it is beneficial to ride in a more upright TT position. The two model outputs were not completely similar, due to the different model approaches. The Metabolic Energy Model could be applied for endurance events, while the Power Output Model is more suitable in sprinting or in variable conditions (wind, undulating course, etc.). It is suggested that despite some limitations, the models give valuable information about improving the cycling performance by optimising the TT cycling position.
Effect of longitudinal ridges on the aerodynamic performance of a leatherback turtle model
NASA Astrophysics Data System (ADS)
Bang, Kyeongtae; Kim, Jooha; Kim, Heesu; Lee, Sang-Im; Choi, Haecheon
2012-11-01
Leatherback sea turtles (Dermochelys coriacea) are known as the fastest swimmer and the deepest diver in the open ocean among marine turtles. Unlike other marine turtles, leatherback sea turtles have five longitudinal ridges on their carapace. To investigate the effect of these longitudinal ridges on the aerodynamic performance of a leatherback turtle model, the experiment is conducted in a wind tunnel at Re = 1.0 × 105 - 1.4 × 106 (including that of real leatherback turtle in cruising condition) based on the model length. We measure the drag and lift forces on the leatherback turtle model with and without longitudinal ridges. The presence of longitudinal ridges increases both the lift and drag forces on the model, but increases the lift-to-drag ratio by 15 - 40%. We also measure the velocity field around the model with and without the ridges using particle image velocimetry. More details will be shown in the presentation. Supported by the NRF program (2011-0028032).
Relevance of aerodynamic modelling for load reduction control strategies of two-bladed wind turbines
NASA Astrophysics Data System (ADS)
Luhmann, B.; Cheng, P. W.
2014-06-01
A new load reduction concept is being developed for the two-bladed prototype of the Skywind 3.5MW wind turbine. Due to transport and installation advantages both offshore and in complex terrain two-bladed turbine designs are potentially more cost-effective than comparable three-bladed configurations. A disadvantage of two-bladed wind turbines is the increased fatigue loading, which is a result of asymmetrically distributed rotor forces. The innovative load reduction concept of the Skywind prototype consists of a combination of cyclic pitch control and tumbling rotor kinematics to mitigate periodic structural loading. Aerodynamic design tools must be able to model correctly the advanced dynamics of the rotor. In this paper the impact of the aerodynamic modelling approach is investigated for critical operational modes of a two-bladed wind turbine. Using a lifting line free wake vortex code (FVM) the physical limitations of the classical blade element momentum theory (BEM) can be evaluated. During regular operation vertical shear and yawed inflow are the main contributors to periodic blade load asymmetry. It is shown that the near wake interaction of the blades under such conditions is not fully captured by the correction models of BEM approach. The differing prediction of local induction causes a high fatigue load uncertainty especially for two-bladed turbines. The implementation of both cyclic pitch control and a tumbling rotor can mitigate the fatigue loading by increasing the aerodynamic and structural damping. The influence of the time and space variant vorticity distribution in the near wake is evaluated in detail for different cyclic pitch control functions and tumble dynamics respectively. It is demonstrated that dynamic inflow as well as wake blade interaction have a significant impact on the calculated blade forces and need to be accounted for by the aerodynamic modelling approach. Aeroelastic simulations are carried out using the high fidelity multi body
Nabawy, Mostafa R. A.; Crowther, William J.
2014-01-01
This paper introduces a generic, transparent and compact model for the evaluation of the aerodynamic performance of insect-like flapping wings in hovering flight. The model is generic in that it can be applied to wings of arbitrary morphology and kinematics without the use of experimental data, is transparent in that the aerodynamic components of the model are linked directly to morphology and kinematics via physical relationships and is compact in the sense that it can be efficiently evaluated for use within a design optimization environment. An important aspect of the model is the method by which translational force coefficients for the aerodynamic model are obtained from first principles; however important insights are also provided for the morphological and kinematic treatments that improve the clarity and efficiency of the overall model. A thorough analysis of the leading-edge suction analogy model is provided and comparison of the aerodynamic model with results from application of the leading-edge suction analogy shows good agreement. The full model is evaluated against experimental data for revolving wings and good agreement is obtained for lift and drag up to 90° incidence. Comparison of the model output with data from computational fluid dynamics studies on a range of different insect species also shows good agreement with predicted weight support ratio and specific power. The validated model is used to evaluate the relative impact of different contributors to the induced power factor for the hoverfly and fruitfly. It is shown that the assumption of an ideal induced power factor (k = 1) for a normal hovering hoverfly leads to a 23% overestimation of the generated force owing to flapping. PMID:24554578
Nabawy, Mostafa R A; Crowther, William J
2014-05-06
This paper introduces a generic, transparent and compact model for the evaluation of the aerodynamic performance of insect-like flapping wings in hovering flight. The model is generic in that it can be applied to wings of arbitrary morphology and kinematics without the use of experimental data, is transparent in that the aerodynamic components of the model are linked directly to morphology and kinematics via physical relationships and is compact in the sense that it can be efficiently evaluated for use within a design optimization environment. An important aspect of the model is the method by which translational force coefficients for the aerodynamic model are obtained from first principles; however important insights are also provided for the morphological and kinematic treatments that improve the clarity and efficiency of the overall model. A thorough analysis of the leading-edge suction analogy model is provided and comparison of the aerodynamic model with results from application of the leading-edge suction analogy shows good agreement. The full model is evaluated against experimental data for revolving wings and good agreement is obtained for lift and drag up to 90° incidence. Comparison of the model output with data from computational fluid dynamics studies on a range of different insect species also shows good agreement with predicted weight support ratio and specific power. The validated model is used to evaluate the relative impact of different contributors to the induced power factor for the hoverfly and fruitfly. It is shown that the assumption of an ideal induced power factor (k = 1) for a normal hovering hoverfly leads to a 23% overestimation of the generated force owing to flapping.
Uncertainty-Based Approach for Dynamic Aerodynamic Data Acquisition and Analysis
NASA Technical Reports Server (NTRS)
Heim, Eugene H. D.; Bandon, Jay M.
2004-01-01
Development of improved modeling methods to provide increased fidelity of flight predictions for aircraft motions during flight in flow regimes with large nonlinearities requires improvements in test techniques for measuring and characterizing wind tunnel data. This paper presents a method for providing a measure of data integrity for static and forced oscillation test techniques. Data integrity is particularly important when attempting to accurately model and predict flight of today s high performance aircraft which are operating in expanded flight envelopes, often maneuvering at high angular rates at high angles-of-attack, even above maximum lift. Current aerodynamic models are inadequate in predicting flight characteristics in the expanded envelope, such as rapid aircraft departures and other unusual motions. Present wind tunnel test methods do not factor changes of flow physics into data acquisition schemes, so in many cases data are obtained over more iterations than required, or insufficient data may be obtained to determine a valid estimate with statistical significance. Additionally, forced oscillation test techniques, one of the primary tools used to develop dynamic models, do not currently provide estimates of the uncertainty of the results during an oscillation cycle. A method to optimize the required number of forced oscillation cycles based on decay of uncertainty gradients and balance tolerances is also presented.
A versatile low-dimensional vortex model for investigating unsteady aerodynamics
NASA Astrophysics Data System (ADS)
Darakananda, Darwin; Eldredge, Jeff D.
2016-11-01
In previous work, we demonstrated a hybrid vortex sheet/point vortex model that captures the non-linear aerodynamics of a plate translating at a high angle of attack. We used vortex sheets to model the shear layers emerging from the plate, and point vortices to capture the effect of the coherent vortex structures. In this work, we introduce modifications that allow the model to work for a larger range of plate kinematics over longer periods of time. First, following the example of Ramesh et al., we relax the Kutta condition at the leading edge and determine vorticity flux based on a suction parameter instead. To prevent the vortex sheet from becoming unstable near the resulting singular edge, we explicitly filter out short-wave disturbances along the sheet while redistributing the sheet's control points. Second, by looking for intersections between the vortex sheets and any repelling Lagrangian coherent structures, the model can detect the formation of new coherent vortices. Trailing portions of the sheets that become dynamically distinct from the shear layers are rolled up into point vortices. We test these modifications on a variety of problems, including pitch-up, impulsive translation at low angles of attack, as well as flow response to pulse actuation near the leading edge. This work has been supported by AFOSR, under award FA9550-14-1-0328.
NASA Technical Reports Server (NTRS)
Larson, R. S.; Nelson, D. P.; Stevens, B. S.
1979-01-01
Five co-annular nozzle models, covering a systematic variation of nozzle geometry, were tested statically over a range of exhaust conditions including inverted velocity profile (IVP) (fan to primary stream velocity ratio 1) and non IVP profiles. Fan nozzle pressure ratio (FNPR) was varied from 1.3 to 4.1 at primary nozzle pressure ratios (PNPR) of 1.53 and 2.0. Fan stream temperatures of 700 K (1260 deg R) and 1089 K(1960 deg R) were tested with primary stream temperatures of 700 K (1260 deg R), 811 K (1460 deg R), and 1089 K (1960 deg R). At fan and primary stream velocities of 610 and 427 m/sec (2000 and 1400 ft/sec), respectively, increasing fan radius ratio from 0.69 to 0.83 reduced peak perceived noise level (PNL) 3 dB, and an increase in primary radius ratio from 0 to 0.81 (fan radius ratio constant at 0.83) reduced peak PNL an additional 1.0 dB. There were no noise reductions at a fan stream velocity of 853 m/sec (2800 ft/sec). Increasing fan radius ratio from 0.69 to 0.83 reduced nozzle thrust coefficient 1.2 to 1.5% at a PNPR of 1.53, and 1.7 to 2.0% at a PNPR of 2.0. The developed acoustic prediction procedure collapsed the existing data with standard deviation varying from + or - 8 dB to + or - 7 dB. The aerodynamic performance prediction procedure collapsed thrust coefficient measurements to within + or - .004 at a FNPR of 4.0 and a PNPR of 2.0.
A flight-test methodology for identification of an aerodynamic model for a V/STOL aircraft
NASA Technical Reports Server (NTRS)
Bach, R. E., Jr.; Mcnally, B. D.
1989-01-01
This paper describes a flight-test methodology for developing a data base to be used to identify an aerodynamic model of a V/STOL fighter aircraft. The aircraft serves as a test bed at NASA Ames for ongoing research in advanced V/STOL control and display concepts. The flight envelope to be modeled includes hover, transition to conventional flight and back to hover, STOL operation, and normal cruise. Although the aerodynamic model is highly nonlinear, it has been formulated to be linear in the parameters to be identified. Motivation for the flight-test methodology advocated in this paper is based on the choice of a linear least-squares method for model identification. The paper covers elements of the methodology from maneuver design to the completed data base. Major emphasis is placed on the use of state estimation with tracking data to ensure consistency among maneuver variables prior to their entry into the data base. The design and processing of a typical maneuver are illustrated.
A flight-test methodology for identification of an aerodynamic model for a V/STOL aircraft
NASA Technical Reports Server (NTRS)
Bach, Ralph E., Jr.; Mcnally, B. David
1988-01-01
Described is a flight test methodology for developing a data base to be used to identify an aerodynamic model of a vertical and short takeoff and landing (V/STOL) fighter aircraft. The aircraft serves as a test bed at Ames for ongoing research in advanced V/STOL control and display concepts. The flight envelope to be modeled includes hover, transition to conventional flight, and back to hover, STOL operation, and normaL cruise. Although the aerodynamic model is highly nonlinear, it has been formulated to be linear in the parameters to be identified. Motivation for the flight test methodology advocated in this paper is based on the choice of a linear least-squares method for model identification. The paper covers elements of the methodology from maneuver design to the completed data base. Major emphasis is placed on the use of state estimation with tracking data to ensure consistency among maneuver variables prior to their entry into the data base. The design and processing of a typical maneuver is illustrated.
Current Trends in Modeling Research for Turbulent Aerodynamic Flows
NASA Technical Reports Server (NTRS)
Gatski, Thomas B.; Rumsey, Christopher L.; Manceau, Remi
2007-01-01
The engineering tools of choice for the computation of practical engineering flows have begun to migrate from those based on the traditional Reynolds-averaged Navier-Stokes approach to methodologies capable, in theory if not in practice, of accurately predicting some instantaneous scales of motion in the flow. The migration has largely been driven by both the success of Reynolds-averaged methods over a wide variety of flows as well as the inherent limitations of the method itself. Practitioners, emboldened by their ability to predict a wide-variety of statistically steady, equilibrium turbulent flows, have now turned their attention to flow control and non-equilibrium flows, that is, separation control. This review gives some current priorities in traditional Reynolds-averaged modeling research as well as some methodologies being applied to a new class of turbulent flow control problems.
An aerodynamic model for one and two degree of freedom wing rock of slender delta wings
NASA Technical Reports Server (NTRS)
Hong, John
1993-01-01
The unsteady aerodynamic effects due to the separated flow around slender delta wings in motion were analyzed. By combining the unsteady flow field solution with the rigid body Euler equations of motion, self-induced wing rock motion is simulated. The aerodynamic model successfully captures the qualitative characteristics of wing rock observed in experiments. For the one degree of freedom in roll case, the model is used to look into the mechanisms of wing rock and to investigate the effects of various parameters, like angle of attack, yaw angle, displacement of the separation point, and wing inertia. To investigate the roll and yaw coupling for the delta wing, an additional degree of freedom is added. However, no limit cycle was observed in the two degree of freedom case. Nonetheless, the model can be used to apply various control laws to actively control wing rock using, for example, the displacement of the leading edge vortex separation point by inboard span wise blowing.
Development of the Dual Aerodynamic Nozzle Model for the NTF Semi-Span Model Support System
NASA Technical Reports Server (NTRS)
Jones, Greg S.; Milholen, William E., II; Goodliff, Scott L.
2011-01-01
The recent addition of a dual flow air delivery system to the NASA Langley National Transonic Facility was experimentally validated with a Dual Aerodynamic Nozzle semi-span model. This model utilized two Stratford calibration nozzles to characterize the weight flow system of the air delivery system. The weight flow boundaries for the air delivery system were identified at mildly cryogenic conditions to be 0.1 to 23 lbm/sec for the high flow leg and 0.1 to 9 lbm/sec for the low flow leg. Results from this test verified system performance and identified problems with the weight-flow metering system that required the vortex flow meters to be replaced at the end of the test.
Identification of an unsteady aerodynamic model up to high angle of attack regime
NASA Astrophysics Data System (ADS)
Fan, Yigang
1997-12-01
The harmonic oscillatory tests for a fighter aircraft configuration using the Dynamic Plunge-Pitch-Roll (DyPPiR) model mount at Virginia Tech Stability Wind Tunnel are described and analyzed. The corresponding data reduction methods are developed on the basis of multirate digital signal processing techniques. Since the model is sting-mounted to the support system of DyPPiR, the Discrete Fourier Transform (DFT) is first used to identify the frequencies of the elastic modes of sting. Then the sampling rate conversion systems are built up in digital domain to resample the data at a lower rate without introducing distortions to the signals of interest. Finally linear-phase Finite Impulse Response (FIR) filters are designed by Remez exchange algorithm to extract the aerodynamic characteristics responses to the programmed motions from the resampled measurements. These data reduction procedures are also illustrated through examples. The results obtained from the harmonic oscillatory tests are then illustrated and the associated flow mechanisms are discussed. Since no significant hysteresis loops are observed for the lift and the drag coefficients for the current angle of attack range and the tested reduced frequencies, the dynamic lags of separated and vortex flow effects are small in the current oscillatory tests. However, large hysteresis loops are observed for pitch moment coefficient in the current tests. This observation suggests that at current flow conditions, pitch moment has large pitch rate dotalpha dependencies. Then the nondimensional maximum pitch rate \\ qsb{max} is introduced to characterize these harmonic oscillatory motions. It is found that at current flow conditions, all the hysteresis loops of pitch moment coefficient with same \\ qsb{max} are tangential to one another at both top and bottom of the loops, implying approximately same maximum offset of these loops from static values. Several cases are also illustrated. Based on the results obtained and
Bird Flight as a Model for a Course in Unsteady Aerodynamics
NASA Astrophysics Data System (ADS)
Jacob, Jamey; Mitchell, Jonathan; Puopolo, Michael
2014-11-01
Traditional unsteady aerodynamics courses at the graduate level focus on theoretical formulations of oscillating airfoil behavior. Aerodynamics students with a vision for understanding bird-flight and small unmanned aircraft dynamics desire to move beyond traditional flow models towards new and creative ways of appreciating the motion of agile flight systems. High-speed videos are used to record kinematics of bird flight, particularly barred owls and red-shouldered hawks during perching maneuvers, and compared with model aircraft performing similar maneuvers. Development of a perching glider and associated control laws to model the dynamics are used as a class project. Observations are used to determine what different species and sizes of birds share in their methods to approach a perch under similar conditions. Using fundamental flight dynamics, simplified models capable of predicting position, attitude, and velocity of the flier are developed and compared with the observations. By comparing the measured data from the videos and predicted and measured motions from the glider models, it is hoped that the students gain a better understanding of the complexity of unsteady aerodynamics and aeronautics and an appreciation for the beauty of avian flight.
Aeroacoustic Study of a High-Fidelity Aircraft Model: Part 1- Steady Aerodynamic Measurements
NASA Technical Reports Server (NTRS)
Khorrami, Mehdi R.; Hannon, Judith A.; Neuhart, Danny H.; Markowski, Gregory A.; VandeVen, Thomas
2012-01-01
In this paper, we present steady aerodynamic measurements for an 18% scale model of a Gulfstream air-craft. The high fidelity and highly-instrumented semi-span model was developed to perform detailed aeroacoustic studies of airframe noise associated with main landing gear/flap components and gear-flap interaction noise, as well as to evaluate novel noise reduction concepts. The aeroacoustic tests, being conducted in the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel, are split into two entries. The first entry, completed November 2010, was entirely devoted to the detailed mapping of the aerodynamic characteristics of the fabricated model. Flap deflections of 39?, 20?, and 0? with the main landing gear on and off were tested at Mach numbers of 0.16, 0.20, and 0.24. Additionally, for each flap deflection, the model was tested with the tunnel both in the closed-wall and open-wall (jet) modes. During this first entry, global forces (lift and drag) and extensive steady and unsteady surface pressure measurements were obtained. Preliminary analysis of the measured forces indicates that lift, drag, and stall characteristics compare favorably with Gulfstream?s high Reynolds number flight data. The favorable comparison between wind-tunnel and flight data allows the semi-span model to be used as a test bed for developing/evaluating airframe noise reduction concepts under a relevant environment. Moreover, initial comparison of the aerodynamic measurements obtained with the tunnel in the closed- and open-wall configurations shows similar aerodynamic behavior. This permits the acoustic and off-surface flow measurements, planned for the second entry, to be conducted with the tunnel in the open-jet mode.
Aerodynamic wake study: oscillating model wind turbine within a turbulent boundary layer
NASA Astrophysics Data System (ADS)
Feist, Christopher J.
An experimental investigation on the aerodynamic wake behind a pitching and/or heaving model wind turbine was performed. The study was split into two quasi-coupled phases; the first phase characterized the motion of an offshore floating wind turbine subjected to linear wave forcing, the second phase replicated specific motion cases, which were driven by results from phase I, on a model wind turbine within a turbulent boundary layer. Wake measurements were made in an effort to quantify fluctuations in the flow associated with the motion of the turbine. Weak differences were observed in the mean, streamwise velocity and turbulent fluctuations between the static and oscillating turbine cases. These weak differences were a result of opposing trends in the velocity quantities based on turbine motion phases. The wake oscillations created by the turbine motion was characteristic of a 2D wave (with convection in the x plane and amplitude in the z plane) with a relatively small amplitude as compared to urms..
NASA Technical Reports Server (NTRS)
Tiffany, Sherwood H.; Karpel, Mordechay
1989-01-01
Various control analysis, design, and simulation techniques for aeroelastic applications require the equations of motion to be cast in a linear time-invariant state-space form. Unsteady aerodynamics forces have to be approximated as rational functions of the Laplace variable in order to put them in this framework. For the minimum-state method, the number of denominator roots in the rational approximation. Results are shown of applying various approximation enhancements (including optimization, frequency dependent weighting of the tabular data, and constraint selection) with the minimum-state formulation to the active flexible wing wind-tunnel model. The results demonstrate that good models can be developed which have an order of magnitude fewer augmenting aerodynamic equations more than traditional approaches. This reduction facilitates the design of lower order control systems, analysis of control system performance, and near real-time simulation of aeroservoelastic phenomena.
Unsteady Aerodynamic Models for Turbomachinery Aeroelastic and Aeroacoustic Applications
NASA Technical Reports Server (NTRS)
Verdon, Joseph M.; Barnett, Mark; Ayer, Timothy C.
1995-01-01
Theoretical analyses and computer codes are being developed for predicting compressible unsteady inviscid and viscous flows through blade rows of axial-flow turbomachines. Such analyses are needed to determine the impact of unsteady flow phenomena on the structural durability and noise generation characteristics of the blading. The emphasis has been placed on developing analyses based on asymptotic representations of unsteady flow phenomena. Thus, high Reynolds number flows driven by small amplitude unsteady excitations have been considered. The resulting analyses should apply in many practical situations and lead to a better understanding of the relevant flow physics. In addition, they will be efficient computationally, and therefore, appropriate for use in aeroelastic and aeroacoustic design studies. Under the present effort, inviscid interaction and linearized inviscid unsteady flow models have been formulated, and inviscid and viscid prediction capabilities for subsonic steady and unsteady cascade flows have been developed. In this report, we describe the linearized inviscid unsteady analysis, LINFLO, the steady inviscid/viscid interaction analysis, SFLOW-IVI, and the unsteady viscous layer analysis, UNSVIS. These analyses are demonstrated via application to unsteady flows through compressor and turbine cascades that are excited by prescribed vortical and acoustic excitations and by prescribed blade vibrations. Recommendations are also given for the future research needed for extending and improving the foregoing asymptotic analyses, and to meet the goal of providing efficient inviscid/viscid interaction capabilities for subsonic and transonic unsteady cascade flows.
Neural Net-Based Redesign of Transonic Turbines for Improved Unsteady Aerodynamic Performance
NASA Technical Reports Server (NTRS)
Madavan, Nateri K.; Rai, Man Mohan; Huber, Frank W.
1998-01-01
A recently developed neural net-based aerodynamic design procedure is used in the redesign of a transonic turbine stage to improve its unsteady aerodynamic performance. The redesign procedure used incorporates the advantages of both traditional response surface methodology (RSM) and neural networks by employing a strategy called parameter-based partitioning of the design space. Starting from the reference design, a sequence of response surfaces based on both neural networks and polynomial fits are constructed to traverse the design space in search of an optimal solution that exhibits improved unsteady performance. The procedure combines the power of neural networks and the economy of low-order polynomials (in terms of number of simulations required and network training requirements). A time-accurate, two-dimensional, Navier-Stokes solver is used to evaluate the various intermediate designs and provide inputs to the optimization procedure. The optimization procedure yields a modified design that improves the aerodynamic performance through small changes to the reference design geometry. The computed results demonstrate the capabilities of the neural net-based design procedure, and also show the tremendous advantages that can be gained by including high-fidelity unsteady simulations that capture the relevant flow physics in the design optimization process.
Aerodynamic characteristics of the standard dynamics model in coning motion at Mach 0.6
NASA Technical Reports Server (NTRS)
Jermey, C.; Schiff, L. B.
1985-01-01
A wind tunnel test was conducted on the Standard Dynamics Model (a simplified generic fighter aircraft shape) undergoing coning motion at Mach 0.6. Six component force and moment data are presented for a range of angle of attack, sideslip, and coning rates. At the relatively low non-dimensional coning rate employed (omega b/2V less than or equal to 0.04), the lateral aerodynamic characteristics generally show a linear variation with coning rate.
A Comprehensive Analytical Model of Rotorcraft Aerodynamics and Dynamics. Part 2. User’s Manual
1980-07-01
Laboratories SUMMARY The use of a comprehensive analytical model of rotorcraft aerodynamics and dynamics is described. This analysis is designed to calculate ...computer program calculates the loads and motion of helicopter rotors and airframe. First the trim soilution is obtained; then the flutter, flight...dynamics, and/or transient behavior can be calculated . Either a new job can be initiated, or further calculations can be performed for an old job. For a new
Gromke, Christof
2011-01-01
A new vegetation modeling concept for Building and Environmental Aerodynamics wind tunnel investigations was developed. The modeling concept is based on fluid dynamical similarity aspects and allows the small-scale modeling of various kinds of vegetation, e.g. field crops, shrubs, hedges, single trees and forest stands. The applicability of the modeling concept was validated in wind tunnel pollutant dispersion studies. Avenue trees in urban street canyons were modeled and their implications on traffic pollutant dispersion were investigated. The dispersion experiments proved the modeling concept to be practicable for wind tunnel studies and suggested to provide reliable concentration results. Unfavorable effects of trees on pollutant dispersion and natural ventilation in street canyons were revealed. Increased traffic pollutant concentrations were found in comparison to the tree-free reference case.
Comparison of aerodynamic models for Vertical Axis Wind Turbines
NASA Astrophysics Data System (ADS)
Simão Ferreira, C.; Aagaard Madsen, H.; Barone, M.; Roscher, B.; Deglaire, P.; Arduin, I.
2014-06-01
Multi-megawatt Vertical Axis Wind Turbines (VAWTs) are experiencing an increased interest for floating offshore applications. However, VAWT development is hindered by the lack of fast, accurate and validated simulation models. This work compares six different numerical models for VAWTS: a multiple streamtube model, a double-multiple streamtube model, the actuator cylinder model, a 2D potential flow panel model, a 3D unsteady lifting line model, and a 2D conformal mapping unsteady vortex model. The comparison covers rotor configurations with two NACA0015 blades, for several tip speed ratios, rotor solidity and fixed pitch angle, included heavily loaded rotors, in inviscid flow. The results show that the streamtube models are inaccurate, and that correct predictions of rotor power and rotor thrust are an effect of error cancellation which only occurs at specific configurations. The other four models, which explicitly model the wake as a system of vorticity, show mostly differences due to the instantaneous or time averaged formulation of the loading and flow, for which further research is needed.
NASA Technical Reports Server (NTRS)
Thorp, Scott A.
1992-01-01
This presentation will discuss the development of a NASA Geometry Exchange Specification for transferring aerodynamic surface geometry between LeRC systems and grid generation software used for computational fluid dynamics research. The proposed specification is based on a subset of the Initial Graphics Exchange Specification (IGES). The presentation will include discussion of how the NASA-IGES standard will accommodate improved computer aided design inspection methods and reverse engineering techniques currently being developed. The presentation is in viewgraph format.
Uncovering the aerodynamics of the smallest insects using numerical and physical models
NASA Astrophysics Data System (ADS)
Miller, Laura
2011-11-01
A vast body of research has described the complexity of flight in insects ranging from the fruit fly, Drosophila melanogaster, to the hawk moth, Manduca sexta. The smallest flying insects have received far less attention, although previous work has shown that flight kinematics and aerodynamics can be significantly different. In this presentation, three-dimensional direct numerical simulations are used to compute the lift and drag forces generated by flexible wings to reveal the aerodynamics of these tiny fliers. Results are validated against dynamically scaled physical models. At the lowest Reynolds numbers relevant to insect flight, the relative forces required to rotate the wings and fling them apart become substantially greater. Wing flexibility can reduce these forces and improve efficiency in some situations.
Optical Model Reduction and Robust Feedback Control for Aerodynamics
2010-03-29
the distance minimization (7) is then the minimal distance from T to Hilbert-Schmidt operators of rank n. In other terms, we have min HSs S T s...coefficien chord (su A The NAC rolled flow at the flow t Model E Surface S ction, we co r is a model is
NASA Astrophysics Data System (ADS)
van Dam, Cornelis P.; Nakafuji, Dora Y.; Bauer, Candice; Standish, Kevin; Chao, David
2003-01-01
A computational design and analysis of a microtab based aerodynamic loads control system is presented. The microtab consists of a small tab that emerges from a wing approximately perpendicular to its surface in the vicinity of its trailing edge. Tab deployment on the upper side of the wing causes a decrease in the lift generation whereas deployment on the pressure side causes an increase. The computational methods applied in the development of this concept solve the governing Reynolds-averaged Navier-Stokes equations on structured, overset grids. The application of these methods to simulate the flows over lifting surfaces including the tabs has been paramount in the development of these devices. The numerical results demonstrate the effectiveness of the microtab and that it is possible to carry out a sensitivity analysis on the positioning and sizing of the tabs before they are implemented in successfully controlling the aerodynamic loads.
NASA Technical Reports Server (NTRS)
Spearman, M. L.; Fournier, R. H.; Lamb, M.
1977-01-01
The aerodynamic, stability, and control characteristics of several supersonic fighter airplane concepts are examined. The configurations, which are based on Soviet design concepts, include fixed-wing aircraft having delta wings, swept wings, and trapezoidal wings, and a variable wing-sweep aircraft. Each concept employs aft tail controls. The concepts vary from lightweight, single-engine, air superiority, point interceptor, or ground attack types to larger twin-engine interceptor and reconnaissance designs. Analytical and experimental results indicate that careful application of the transonic or supersonic area rule can provide nearly optimum shaping for minimum drag for a specified Mach number requirement. In addition, through the proper location of components and the exploitation of interference flow fields, the concepts provide linear pitching moment characteristics, high control effectiveness, and reasonably small variations in aerodynamic center location with a resulting high potential for maneuvering capability.
NASA Technical Reports Server (NTRS)
Hoad, D. R.
1974-01-01
An investigation of a four-engine externally blown flap (EBF) powered-lift transport was conducted in the Langley V/STOL tunnel to determine the effect of different engine configurations on the longitudinal aerodynamic characteristics. The different engine configurations were simulated by five different sets of propulsion simulators on a single aircraft model. Longitudinal aerodynamic data were obtained for each simulator on each flap deflection corresponding to cruise, take-off, and landing at a range of angles of attack and various thrust coefficients. The bypass ratio (BPR) 6.2 engine simulator provided the best lift and drag characteristics of the five simulators tested in the take-off and landing configurations. The poor performance of the BPR 10.0 and 3.2 engine simulators can be attributed to a mismatch of engine-model sizes or poor engine location and orientation. Isolated engine wake surveys indicated that a reasonable assessment of the aerodynamic characteristics of an engine-wing-flap configuration could be made if qualitative information were available which defined the engine wake characteristics. All configurations could be trimmed easily with relatively small horizontal-tail incidence angles; however, the take-off landing configurations required a high-lift tail.
Aerodynamic and base heating studies on space shuttle configurations
NASA Technical Reports Server (NTRS)
1974-01-01
Heating rate and pressure measurements were obtained on a 25-O space shuttle model in a vacuum chamber. Correlation data on windward laminar and turbulent boundary layers and leeside surfaces of the space shuttle orbiter are included.
Turbulence Modeling and Computation of Turbine Aerodynamics and Heat Transfer
NASA Technical Reports Server (NTRS)
Lakshminarayana, B.; Luo, J.
1996-01-01
The objective of the present research is to develop improved turbulence models for the computation of complex flows through turbomachinery passages, including the effects of streamline curvature, heat transfer and secondary flows. Advanced turbulence models are crucial for accurate prediction of rocket engine flows, due to existance of very large extra strain rates, such as strong streamline curvature. Numerical simulation of the turbulent flows in strongly curved ducts, including two 180-deg ducts, one 90-deg duct and a strongly concave curved turbulent boundary layer have been carried out with Reynolds stress models (RSM) and algebraic Reynolds stress models (ARSM). An improved near-wall pressure-strain correlation has been developed for capturing the anisotropy of turbulence in the concave region. A comparative study of two modes of transition in gas turbine, the by-pass transition and the separation-induced transition, has been carried out with several representative low-Reynolds number (LRN) k-epsilon models. Effects of blade surface pressure gradient, freestream turbulence and Reynolds number on the blade boundary layer development, and particularly the inception of transition are examined in detail. The present study indicates that the turbine blade transition, in the presence of high freestream turbulence, is predicted well with LRN k-epsilon models employed. The three-dimensional Navier-Stokes procedure developed by the present authors has been used to compute the three-dimensional viscous flow through the turbine nozzle passage of a single stage turbine. A low Reynolds number k-epsilon model and a zonal k-epsilon/ARSM (algebraic Reynolds stress model) are utilized for turbulence closure. An assessment of the performance of the turbulence models has been carried out. The two models are found to provide similar predictions for the mean flow parameters, although slight improvement in the prediction of some secondary flow quantities has been obtained by the
Numerical modeling of wind turbine aerodynamic noise in the time domain.
Lee, Seunghoon; Lee, Seungmin; Lee, Soogab
2013-02-01
Aerodynamic noise from a wind turbine is numerically modeled in the time domain. An analytic trailing edge noise model is used to determine the unsteady pressure on the blade surface. The far-field noise due to the unsteady pressure is calculated using the acoustic analogy theory. By using a strip theory approach, the two-dimensional noise model is applied to rotating wind turbine blades. The numerical results indicate that, although the operating and atmospheric conditions are identical, the acoustical characteristics of wind turbine noise can be quite different with respect to the distance and direction from the wind turbine.
Calibrated Blade-Element/Momentum Theory Aerodynamic Model of the MARIN Stock Wind Turbine: Preprint
Goupee, A.; Kimball, R.; de Ridder, E. J.; Helder, J.; Robertson, A.; Jonkman, J.
2015-04-02
In this paper, a calibrated blade-element/momentum theory aerodynamic model of the MARIN stock wind turbine is developed and documented. The model is created using open-source software and calibrated to closely emulate experimental data obtained by the DeepCwind Consortium using a genetic algorithm optimization routine. The provided model will be useful for those interested in validating interested in validating floating wind turbine numerical simulators that rely on experiments utilizing the MARIN stock wind turbine—for example, the International Energy Agency Wind Task 30’s Offshore Code Comparison Collaboration Continued, with Correlation project.
NASA Technical Reports Server (NTRS)
Trescot, C. D., Jr.; Brown, C. A., Jr.; Howell, D. T.
1974-01-01
An investigation has been made in the Langley Unitary Plan wind tunnel to determine the effects of Reynolds number and sting-support interference on the static aerodynamic characteristics of a 140 deg-included-angle cone. Base pressures and forces and moments of the model were measured at Mach numbers of 1.50, 2.00, 2.94, and 4.00 for ratios of sting diameter to model diameter that varied from 0.125 to 0.500 through an angle-of-attack range from about minus 4 deg to 13 deg. The Reynolds number, based on model diameter 4.80 in. was varied from 161,000 to 415,000.
Esophageal aerodynamics in an idealized experimental model of tracheoesophageal speech
NASA Astrophysics Data System (ADS)
Erath, Byron D.; Hemsing, Frank S.
2016-03-01
Flow behavior is investigated in the esophageal tract in an idealized experimental model of tracheoesophageal speech. The tracheoesophageal prosthesis is idealized as a first-order approximation using a straight, constant diameter tube. The flow is scaled according to Reynolds, Strouhal, and Euler numbers to ensure dynamic similarity. Flow pulsatility is produced by a driven orifice that approximates the kinematics of the pharyngoesophageal segment during tracheoesophageal speech. Particle image velocimetry data are acquired in three orthogonal planes as the flow exits the model prosthesis and enters the esophageal tract. Contrary to prior investigations performed in steady flow with the prosthesis oriented in-line with the flow direction, the fluid dynamics are shown to be highly unsteady, suggesting that the esophageal pressure field will be similarly complex. A large vortex ring is formed at the inception of each phonatory cycle, followed by the formation of a persistent jet. This vortex ring appears to remain throughout the entire cycle due to the continued production of vorticity resulting from entrainment between the prosthesis jet and the curved esophageal walls. Mean flow in the axial direction of the esophagus produces significant stretching of the vortex throughout the phonatory cycle. The stagnation point created by the jet impinging on the esophageal wall varies throughout the cycle due to fluctuations in the jet trajectory, which most likely arises due to flow separation within the model prosthesis. Applications to tracheoesophageal speech, including shortcomings of the model and proposed future plans, are discussed.
Non-Equilibrium Turbulence Modeling for High Lift Aerodynamics
NASA Technical Reports Server (NTRS)
Durbin, P. A.
1998-01-01
This phase is discussed in ('Non linear kappa - epsilon - upsilon(sup 2) modeling with application to high lift', Application of the kappa - epsilon -upsilon(sup 2) model to multi-component airfoils'). Further results are presented in 'Non-linear upsilon(sup 2) - f modeling with application to high-lift' The ADI solution method in the initial implementation was very slow to converge on multi-zone chimera meshes. I modified the INS implementation to use GMRES. This provided improved convergence and less need for user intervention in the solution process. There were some difficulties with implementation into the NASA compressible codes, due to their use of approximate factorization. The Helmholtz equation for f is not an evolution equation, so it is not of the form assumed by the approximate factorization method. Although The Kalitzin implementation involved a new solution algorithm ('An implementation of the upsilon(sup 2) - f model with application to transonic flows'). The algorithm involves introducing a relaxation term in the f-equation so that it can be factored. The factorization can be into a plane and a line, with GMRES used in the plane. The NASA code already evaluated coefficients in planes, so no additional memory is required except that associated the the GMRES algorithm. So the scope of this project has expanded via these interactions. . The high-lift work has dovetailed into turbine applications.
NASA Technical Reports Server (NTRS)
Wittliff, C. E.
1982-01-01
The aerodynamic heating of a tip-fin controller mounted on a Space Shuttle Orbiter model was studied experimentally in the Calspan Advanced Technology Center 96 inch Hypersonic Shock Tunnel. A 0.0175 scale model was tested at Mach numbers from 10 to 17.5 at angles of attack typical of a shuttle entry. The study was conducted in two phases. In phase 1 testing a thermographic phosphor technique was used to qualitatively determine the areas of high heat-transfer rates. Based on the results of this phase, the model was instrumented with 40 thin-film resistance thermometers to obtain quantitative measurements of the aerodynamic heating. The results of the phase 2 testing indicate that the highest heating rates, which occur on the leading edge of the tip-fin controller, are very sensitive to angle of attack for alpha or = 30 deg. The shock wave from the leading edge of the orbiter wing impinges on the leading edge of the tip-fin controller resulting in peak values of h/h(Ref) in the range from 1.5 to 2.0. Away from the leading edge, the heat-transfer rates never exceed h/h(Ref) = 0.25 when the control surface, is not deflected. With the control surface deflected 20 deg, the heat-transfer rates had a maximum value of h/h(Ref) = 0.3. The heating rates are quite nonuniform over the outboard surface and are sensitive to angle of attack.
Aerodynamic flight evaluation analysis and data base update
NASA Technical Reports Server (NTRS)
Boyle, W. W.; Miller, M. S.; Wilder, G. O.; Reheuser, R. D.; Sharp, R. S.; Bridges, G. I.
1989-01-01
Research was conducted to determine the feasibility of replacing the Solid Rocket Boosters on the existing Space Shuttle Launch Vehicle (SSLV) with Liquid Rocket Boosters (LRB). As a part of the LRB selection process, a series of wind tunnel tests were conducted along with aero studies to determine the effects of different LRB configurations on the SSLV. Final results were tabulated into increments and added to the existing SSLV data base. The research conducted in this study was taken from a series of wind tunnel tests conducted at Marshall's 14-inch Trisonic Wind Tunnel. The effects on the axial force (CAF), normal force (CNF), pitching moment (CMF), side force (CY), wing shear force (CSR), wing torque moment (CTR), and wing bending moment (CBR) coefficients were investigated for a number of candidate LRB configurations. The aero effects due to LRB protuberances, ET/LRB separation distance, and aft skirts were also gathered from the tests. Analysis was also conducted to investigate the base pressure and plume effects due to the new booster geometries. The test results found in Phases 1 and 2 of wind tunnel testing are discussed and compared. Preliminary LRB lateral/directional data results and trends are given. The protuberance and gap/skirt effects are discussed. The base pressure/plume effects study is discussed and results are given.
NASA Technical Reports Server (NTRS)
Ferris, J. C.
1986-01-01
A wind-tunnel investigation was made to determine the longitudinal aerodynamic characteristics of a fixed-wing generic fighter model with a wing designed for sustained transonic maneuver conditions. The airfoil sections on the wing were designed with a two-dimensional nonlinear computer code, and the root and tip section were modified with a three-dimensional code. The wing geometric characteristics were as follows: a leading-edge sweep of 45 degrees, a taper ratio of 0.2141, an aspect ratio of 3.30, and a thickness ratio of 0.044. The model was investigated at Mach numbers from 0.600 to 1.200, at Reynolds numbers, based on the model reference length, from 2,560,000 to 3,970,000, and through a model angle-of-attack range from -5 to +18 degrees.
Modelling of plasma aerodynamic actuation driven by nanosecond SDBD discharge
NASA Astrophysics Data System (ADS)
Zhu, Yifei; Wu, Yun; Cui, Wei; Li, Yinghong; Jia, Min
2013-09-01
A two-dimensional air plasma kinetics model (16 species and 44 processes) for nanosecond discharge under atmospheric pressure was developed to reveal the spatial and temporal distribution of discharge characteristics of a surface dielectric barrier discharge (SDBD) actuator. An energy transfer model, including two channels for energy release from external power source to gas, was developed to couple plasma with hydrodynamics directly in the same dimension. The governing equations included the Poisson equation for the electric potential, continuity equations for each species, electron energy equations for electrons taking part in reactions, and Navier-Stokes equations for non-isothermal fluid. The model was validated through current-voltage profile and electron temperature obtained from experiments. Calculations for discharge characteristics as well as the responses of fluid field from tens of nanoseconds to tens of seconds were performed. Results have shown that local air is heated to 1170 K within tens of nanoseconds and then decreases to 310 K at the end of a discharge period. 30% of the total power is transferred from electric field to electrons while only 20% of this energy is then released to gas through quenching processes. 9% of the total energy is released through ion collision. A micro-shock wave is formed and propagates at the speed of sound. High local density gradient and dynamic viscosity induces vortexes which whirl the heated air downstream. The combined effects of heating convection and vortexes in repetitive pulse discharges lead to the formation of a steady jet, in agreement with experimental results.
Aerodynamics on a transport aircraft type wing-body model
NASA Technical Reports Server (NTRS)
Schmitt, V.
1982-01-01
The DFLR-F4 wing-body combination is studied. The 1/38 model is formed by a 9.5 aspect ratio transonic wing and an Airbus A 310 fuselage. The F4 wing geometrical characteristics are described and the main experimental results obtained in the S2MA wind tunnel are discussed. Both wing-fuselage interferences and viscous effects, which are important on the wing due to a high rear loading, are investigated by performing 3D calculations. An attempt is made to find their limitations.
Effective Inflow Conditions for Turbulence Models in Aerodynamic Calculations
NASA Technical Reports Server (NTRS)
Spalart, Philippe R.; Rumsey, Christopher L.
2007-01-01
The selection of inflow values at boundaries far upstream of an aircraft is considered, for one- and two-equation turbulence models. Inflow values are distinguished from the ambient values near the aircraft, which may be much smaller. Ambient values should be selected first, and inflow values that will lead to them after the decay second; this is not always possible, especially for the time scale. The two-equation decay during the approach to the aircraft is shown; often, the time scale has been set too short for this decay to be calculated accurately on typical grids. A simple remedy for both issues is to impose floor values for the turbulence variables, outside the viscous sublayer, and it is argued that overriding the equations in this manner is physically justified. Selecting laminar ambient values is easy, if the boundary layers are to be tripped, but a more common practice is to seek ambient values that will cause immediate transition in boundary layers. This opens up a wide range of values, and selection criteria are discussed. The turbulent Reynolds number, or ratio of eddy viscosity to laminar viscosity has a huge dynamic range that makes it unwieldy; it has been widely mis-used, particularly by codes that set upper limits on it. The value of turbulent kinetic energy in a wind tunnel or the atmosphere is also of dubious value as an input to the model. Concretely, the ambient eddy viscosity must be small enough to preserve potential cores in small geometry features, such as flap gaps. The ambient frequency scale should also be small enough, compared with shear rates in the boundary layer. Specific values are recommended and demonstrated for airfoil flows
Gradient-based optimum aerodynamic design using adjoint methods
NASA Astrophysics Data System (ADS)
Xie, Lei
2002-09-01
Continuous adjoint methods and optimal control theory are applied to a pressure-matching inverse design problem of quasi 1-D nozzle flows. Pontryagin's Minimum Principle is used to derive the adjoint system and the reduced gradient of the cost functional. The properties of adjoint variables at the sonic throat and the shock location are studied, revealing a log-arithmic singularity at the sonic throat and continuity at the shock location. A numerical method, based on the Steger-Warming flux-vector-splitting scheme, is proposed to solve the adjoint equations. This scheme can finely resolve the singularity at the sonic throat. A non-uniform grid, with points clustered near the throat region, can resolve it even better. The analytical solutions to the adjoint equations are also constructed via Green's function approach for the purpose of comparing the numerical results. The pressure-matching inverse design is then conducted for a nozzle parameterized by a single geometric parameter. In the second part, the adjoint methods are applied to the problem of minimizing drag coefficient, at fixed lift coefficient, for 2-D transonic airfoil flows. Reduced gradients of several functionals are derived through application of a Lagrange Multiplier Theorem. The adjoint system is carefully studied including the adjoint characteristic boundary conditions at the far-field boundary. A super-reduced design formulation is also explored by treating the angle of attack as an additional state; super-reduced gradients can be constructed either by solving adjoint equations with non-local boundary conditions or by a direct Lagrange multiplier method. In this way, the constrained optimization reduces to an unconstrained design problem. Numerical methods based on Jameson's finite volume scheme are employed to solve the adjoint equations. The same grid system generated from an efficient hyperbolic grid generator are adopted in both the Euler flow solver and the adjoint solver. Several
Aerodynamic Measurements of a Gulfstream Aircraft Model With and Without Noise Reduction Concepts
NASA Technical Reports Server (NTRS)
Neuhart, Dan H.; Hannon, Judith A.; Khorrami, Mehdi R.
2014-01-01
Steady and unsteady aerodynamic measurements of a high-fidelity, semi-span 18% scale Gulfstream aircraft model are presented. The aerodynamic data were collected concurrently with acoustic measurements as part of a larger aeroacoustic study targeting airframe noise associated with main landing gear/flap components, gear-flap interaction noise, and the viability of related noise mitigation technologies. The aeroacoustic tests were conducted in the NASA Langley Research Center 14- by 22-Foot Subsonic Wind Tunnel with the facility in the acoustically treated open-wall (jet) mode. Most of the measurements were obtained with the model in landing configuration with the flap deflected at 39º and the main landing gear on and off. Data were acquired at Mach numbers of 0.16, 0.20, and 0.24. Global forces (lift and drag) and extensive steady and unsteady surface pressure measurements were obtained. Comparison of the present results with those acquired during a previous test shows a significant reduction in the lift experienced by the model. The underlying cause was traced to the likely presence of a much thicker boundary layer on the tunnel floor, which was acoustically treated for the present test. The steady and unsteady pressure fields on the flap, particularly in the regions of predominant noise sources such as the inboard and outboard tips, remained unaffected. It is shown that the changes in lift and drag coefficients for model configurations fitted with gear/flap noise abatement technologies fall within the repeatability of the baseline configuration. Therefore, the noise abatement technologies evaluated in this experiment have no detrimental impact on the aerodynamic performance of the aircraft model.
Analytical model of rotor wake aerodynamics in ground effect
NASA Technical Reports Server (NTRS)
Saberi, H. A.
1983-01-01
The model and the computer program developed provides the velocity, location, and circulation of the tip vortices of a two-blade helicopter in and out of the ground effect. Comparison of the theoretical results with some experimental measurements for the location of the wake indicate that there is excellent accuracy in the vicinity of the rotor and fair amount of accuracy far from it. Having the location of the wake at all times enables us to compute the history of the velocity and the location of any point in the flow. The main goal of out study, induced velocity at the rotor, can also be calculated in addition to stream lines and streak lines. Since the wake location close to the rotor is known more accurately than at other places, the calculated induced velocity over the disc should be a good estimate of the real induced velocity, with the exception of the blade location, because each blade was replaced only by a vortex line. Because no experimental measurements of the wake close to the ground were available to us, quantitative evaluation of the theoretical wake was not possible. But qualitatively we have been able to show excellent agreement. Comparison of flow visualization with out results has indicated the location of the ground vortex is estimated excellently. Also the flow field in hover is well represented.
Modeling coupled aerodynamics and vocal fold dynamics using immersed boundary methods.
Duncan, Comer; Zhai, Guangnian; Scherer, Ronald
2006-11-01
The penalty immersed boundary (PIB) method, originally introduced by Peskin (1972) to model the function of the mammalian heart, is tested as a fluid-structure interaction model of the closely coupled dynamics of the vocal folds and aerodynamics in phonation. Two-dimensional vocal folds are simulated with material properties chosen to result in self-oscillation and volume flows in physiological frequency ranges. Properties of the glottal flow field, including vorticity, are studied in conjunction with the dynamic vocal fold motion. The results of using the PIB method to model self-oscillating vocal folds for the case of 8 cm H20 as the transglottal pressure gradient are described. The volume flow at 8 cm H20, the transglottal pressure, and vortex dynamics associated with the self-oscillating model are shown. Volume flow is also given for 2, 4, and 12 cm H2O, illustrating the robustness of the model to a range of transglottal pressures. The results indicate that the PIB method applied to modeling phonation has good potential for the study of the interdependence of aerodynamics and vocal fold motion.
NASA Technical Reports Server (NTRS)
Kuhlman, J. M.
1979-01-01
The aerodynamic design of a wind-tunnel model of a wing representative of that of a subsonic jet transport aircraft, fitted with winglets, was performed using two recently developed optimal wing-design computer programs. Both potential flow codes use a vortex lattice representation of the near-field of the aerodynamic surfaces for determination of the required mean camber surfaces for minimum induced drag, and both codes use far-field induced drag minimization procedures to obtain the required spanloads. One code uses a discrete vortex wake model for this far-field drag computation, while the second uses a 2-D advanced panel wake model. Wing camber shapes for the two codes are very similar, but the resulting winglet camber shapes differ widely. Design techniques and considerations for these two wind-tunnel models are detailed, including a description of the necessary modifications of the design geometry to format it for use by a numerically controlled machine for the actual model construction.
Actuator and aerodynamic modeling for high-angle-of-attack aeroservoelasticity
NASA Technical Reports Server (NTRS)
Brenner, Martin J.
1993-01-01
Accurate prediction of airframe/actuation coupling is required by the imposing demands of modern flight control systems. In particular, for agility enhancement at high angle of attack and low dynamic pressure, structural integration characteristics such as hinge moments, effective actuator stiffness, and airframe/control surface damping can have a significant effect on stability predictions. Actuator responses are customarily represented with low-order transfer functions matched to actuator test data, and control surface stiffness is often modeled as a linear spring. The inclusion of the physical properties of actuation and its installation on the airframe is therefore addressed using detailed actuator models which consider the physical, electrical, and mechanical elements of actuation. The aeroservoelastic analysis procedure is described in which the actuators are modeled as detailed high-order transfer functions and as approximate low-order transfer functions. The impacts of unsteady aerodynamic modeling on aeroservoelastic stability are also investigated by varying the order of approximation, or number of aerodynamic lag states, in the analysis. Test data from a thrust-vectoring configuration of an F/A-l8 aircraft are compared to predictions to determine the effects on accuracy as a function of modeling complexity.
Unsteady transonic aerodynamics
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.
Aerodynamic Effects of Simulated Ice Accretion on a Generic Transport Model
NASA Technical Reports Server (NTRS)
Broeren, Andy P.; Lee, Sam; Shah, Gautam H.; Murphy, Patrick C.
2012-01-01
An experimental research effort was begun to develop a database of airplane aerodynamic characteristics with simulated ice accretion over a large range of incidence and sideslip angles. Wind-tunnel testing was performed at the NASA Langley 12-ft Low-Speed Wind Tunnel using a 3.5 percent scale model of the NASA Langley Generic Transport Model. Aerodynamic data were acquired from a six-component force and moment balance in static-model sweeps from alpha = -5deg to 85deg and beta = -45 deg to 45 deg at a Reynolds number of 0.24 x10(exp 6) and Mach number of 0.06. The 3.5 percent scale GTM was tested in both the clean configuration and with full-span artificial ice shapes attached to the leading edges of the wing, horizontal and vertical tail. Aerodynamic results for the clean airplane configuration compared favorably with similar experiments carried out on a 5.5 percent scale GTM. The addition of the large, glaze-horn type ice shapes did result in an increase in airplane drag coefficient but had little effect on the lift and pitching moment. The lateral-directional characteristics showed mixed results with a small effect of the ice shapes observed in some cases. The flow visualization images revealed the presence and evolution of a spanwise-running vortex on the wing that was the dominant feature of the flowfield for both clean and iced configurations. The lack of ice-induced performance and flowfield effects observed in this effort was likely due to Reynolds number effects for the clean configuration. Estimates of full-scale baseline performance were included in this analysis to illustrate the potential icing effects.
Study of aerodynamic structure of flow in a model of vortex furnace using Stereo PIV method
NASA Astrophysics Data System (ADS)
Anufriev, I. S.; Kuibin, P. A.; Shadrin, E. Yu.; Sharaborin, D. K.; Sharypov, O. V.
2016-07-01
The aerodynamic structure of flow in a lab model of a perspective design of vortex furnace was studied. The chamber has a horizontal rotation axis, tangential inlet for fuel-air jets and vertical orientation of secondary injection nozzles. The Stereo PIV method was used for visualization of 3D velocity field for selected cross sections of the vortex combustion chamber. The experimental data along with "total pressure minimum" criterion were used for reconstruction of the vortex core of the flow. Results fit the available data from LDA and simulation.
NASA Technical Reports Server (NTRS)
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).
PREFACE: Aerodynamic sound Aerodynamic sound
NASA Astrophysics Data System (ADS)
Akishita, Sadao
2010-02-01
The modern theory of aerodynamic sound originates from Lighthill's two papers in 1952 and 1954, as is well known. I have heard that Lighthill was motivated in writing the papers by the jet-noise emitted by the newly commercialized jet-engined airplanes at that time. The technology of aerodynamic sound is destined for environmental problems. Therefore the theory should always be applied to newly emerged public nuisances. This issue of Fluid Dynamics Research (FDR) reflects problems of environmental sound in present Japanese technology. The Japanese community studying aerodynamic sound has held an annual symposium since 29 years ago when the late Professor S Kotake and Professor S Kaji of Teikyo University organized the symposium. Most of the Japanese authors in this issue are members of the annual symposium. I should note the contribution of the two professors cited above in establishing the Japanese community of aerodynamic sound research. It is my pleasure to present the publication in this issue of ten papers discussed at the annual symposium. I would like to express many thanks to the Editorial Board of FDR for giving us the chance to contribute these papers. We have a review paper by T Suzuki on the study of jet noise, which continues to be important nowadays, and is expected to reform the theoretical model of generating mechanisms. Professor M S Howe and R S McGowan contribute an analytical paper, a valuable study in today's fluid dynamics research. They apply hydrodynamics to solve the compressible flow generated in the vocal cords of the human body. Experimental study continues to be the main methodology in aerodynamic sound, and it is expected to explore new horizons. H Fujita's study on the Aeolian tone provides a new viewpoint on major, longstanding sound problems. The paper by M Nishimura and T Goto on textile fabrics describes new technology for the effective reduction of bluff-body noise. The paper by T Sueki et al also reports new technology for the
Aerodynamic Performance of Scale-Model Turbofan Outlet Guide Vanes Designed for Low Noise
NASA Technical Reports Server (NTRS)
Hughes, Christopher E.
2001-01-01
The design of effective new technologies to reduce aircraft propulsion noise is dependent on an understanding of the noise sources and noise generation mechanisms in the modern turbofan engine. In order to more fully understand the physics of noise in a turbofan engine, a comprehensive aeroacoustic wind tunnel test programs was conducted called the 'Source Diagnostic Test.' The text was cooperative effort between NASA and General Electric Aircraft Engines, as part of the NASA Advanced Subsonic Technology Noise Reduction Program. A 1/5-scale model simulator representing the bypass stage of a current technology high bypass ratio turbofan engine was used in the test. The test article consisted of the bypass fan and outlet guide vanes in a flight-type nacelle. The fan used was a medium pressure ratio design with 22 individual, wide chord blades. Three outlet guide vane design configurations were investigated, representing a 54-vane radial Baseline configuration, a 26-vane radial, wide chord Low Count configuration and a 26-vane, wide chord Low Noise configuration with 30 deg of aft sweep. The test was conducted in the NASA Glenn Research Center 9 by 15-Foot Low Speed Wind Tunnel at velocities simulating the takeoff and approach phases of the aircraft flight envelope. The Source Diagnostic Test had several acoustic and aerodynamic technical objectives: (1) establish the performance of a scale model fan selected to represent the current technology turbofan product; (2) assess the performance of the fan stage with each of the three distinct outlet guide vane designs; (3) determine the effect of the outlet guide vane configuration on the fan baseline performance; and (4) conduct detailed flowfield diagnostic surveys, both acoustic and aerodynamic, to characterize and understand the noise generation mechanisms in a turbofan engine. This paper addresses the fan and stage aerodynamic performance results from the Source Diagnostic Test.
Evangelista, Dennis; Cardona, Griselda; Guenther-Gleason, Eric; Huynh, Tony; Kwong, Austin; Marks, Dylan; Ray, Neil; Tisbe, Adrian; Tse, Kyle; Koehl, Mimi
2014-01-01
We report the effects of posture and morphology on the static aerodynamic stability and control effectiveness of physical models based on the feathered dinosaur, Microraptor gui, from the Cretaceous of China. Postures had similar lift and drag coefficients and were broadly similar when simplified metrics of gliding were considered, but they exhibited different stability characteristics depending on the position of the legs and the presence of feathers on the legs and the tail. Both stability and the function of appendages in generating maneuvering forces and torques changed as the glide angle or angle of attack were changed. These are significant because they represent an aerial environment that may have shifted during the evolution of directed aerial descent and other aerial behaviors. Certain movements were particularly effective (symmetric movements of the wings and tail in pitch, asymmetric wing movements, some tail movements). Other appendages altered their function from creating yaws at high angle of attack to rolls at low angle of attack, or reversed their function entirely. While M. gui lived after Archaeopteryx and likely represents a side experiment with feathered morphology, the general patterns of stability and control effectiveness suggested from the manipulations of forelimb, hindlimb and tail morphology here may help understand the evolution of flight control aerodynamics in vertebrates. Though these results rest on a single specimen, as further fossils with different morphologies are tested, the findings here could be applied in a phylogenetic context to reveal biomechanical constraints on extinct flyers arising from the need to maneuver. PMID:24454820
Computer graphics in aerodynamic analysis
NASA Technical Reports Server (NTRS)
Cozzolongo, J. V.
1984-01-01
The use of computer graphics and its application to aerodynamic analyses on a routine basis is outlined. The mathematical modelling of the aircraft geometries and the shading technique implemented are discussed. Examples of computer graphics used to display aerodynamic flow field data and aircraft geometries are shown. A future need in computer graphics for aerodynamic analyses is addressed.
Draxl, C.; Churchfield, M.; Mirocha, J.; Lee, S.; Lundquist, J.; Michalakes, J.; Moriarty, P.; Purkayastha, A.; Sprague, M.; Vanderwende, B.
2014-06-01
Wind plant aerodynamics are influenced by a combination of microscale and mesoscale phenomena. Incorporating mesoscale atmospheric forcing (e.g., diurnal cycles and frontal passages) into wind plant simulations can lead to a more accurate representation of microscale flows, aerodynamics, and wind turbine/plant performance. Our goal is to couple a numerical weather prediction model that can represent mesoscale flow [specifically the Weather Research and Forecasting model] with a microscale LES model (OpenFOAM) that can predict microscale turbulence and wake losses.
Vertical Landing Aerodynamics of Reusable Rocket Vehicle
NASA Astrophysics Data System (ADS)
Nonaka, Satoshi; Nishida, Hiroyuki; Kato, Hiroyuki; Ogawa, Hiroyuki; Inatani, Yoshifumi
The aerodynamic characteristics of a vertical landing rocket are affected by its engine plume in the landing phase. The influences of interaction of the engine plume with the freestream around the vehicle on the aerodynamic characteristics are studied experimentally aiming to realize safe landing of the vertical landing rocket. The aerodynamic forces and surface pressure distributions are measured using a scaled model of a reusable rocket vehicle in low-speed wind tunnels. The flow field around the vehicle model is visualized using the particle image velocimetry (PIV) method. Results show that the aerodynamic characteristics, such as the drag force and pitching moment, are strongly affected by the change in the base pressure distributions and reattachment of a separation flow around the vehicle.
Unsteady aerodynamic force generation by a model fruit fly wing in flapping motion.
Sun, Mao; Tang, Jian
2002-01-01
A computational fluid-dynamic analysis was conducted to study the unsteady aerodynamics of a model fruit fly wing. The wing performs an idealized flapping motion that emulates the wing motion of a fruit fly in normal hovering flight. The Navier-Stokes equations are solved numerically. The solution provides the flow and pressure fields, from which the aerodynamic forces and vorticity wake structure are obtained. Insights into the unsteady aerodynamic force generation process are gained from the force and flow-structure information. Considerable lift can be produced when the majority of the wing rotation is conducted near the end of a stroke or wing rotation precedes stroke reversal (rotation advanced), and the mean lift coefficient can be more than twice the quasi-steady value. Three mechanisms are responsible for the large lift: the rapid acceleration of the wing at the beginning of a stroke, the absence of stall during the stroke and the fast pitching-up rotation of the wing near the end of the stroke. When half the wing rotation is conducted near the end of a stroke and half at the beginning of the next stroke (symmetrical rotation), the lift at the beginning and near the end of a stroke becomes smaller because the effects of the first and third mechanisms above are reduced. The mean lift coefficient is smaller than that of the rotation-advanced case, but is still 80 % larger than the quasi-steady value. When the majority of the rotation is delayed until the beginning of the next stroke (rotation delayed), the lift at the beginning and near the end of a stroke becomes very small or even negative because the effect of the first mechanism above is cancelled and the third mechanism does not apply in this case. The mean lift coefficient is much smaller than in the other two cases.
Computational Aerodynamic Analysis of a Micro-CT Based Bio-Realistic Fruit Fly Wing
Brandt, Joshua; Doig, Graham; Tsafnat, Naomi
2015-01-01
The aerodynamic features of a bio-realistic 3D fruit fly wing in steady state (snapshot) flight conditions were analyzed numerically. The wing geometry was created from high resolution micro-computed tomography (micro-CT) of the fruit fly Drosophila virilis. Computational fluid dynamics (CFD) analyses of the wing were conducted at ultra-low Reynolds numbers ranging from 71 to 200, and at angles of attack ranging from -10° to +30°. It was found that in the 3D bio-realistc model, the corrugations of the wing created localized circulation regions in the flow field, most notably at higher angles of attack near the wing tip. Analyses of a simplified flat wing geometry showed higher lift to drag performance values for any given angle of attack at these Reynolds numbers, though very similar performance is noted at -10°. Results have indicated that the simplified flat wing can successfully be used to approximate high-level properties such as aerodynamic coefficients and overall performance trends as well as large flow-field structures. However, local pressure peaks and near-wing flow features induced by the corrugations are unable to be replicated by the simple wing. We therefore recommend that accurate 3D bio-realistic geometries be used when modelling insect wings where such information is useful. PMID:25954946
Flow Quality Measurements in an Aerodynamic Model of NASA Lewis' Icing Research Tunnel
NASA Technical Reports Server (NTRS)
Canacci, Victor A.; Gonsalez, Jose C.
1999-01-01
As part of an ongoing effort to improve the aerodynamic flow characteristics of the Icing Research Tunnel (IRT), a modular scale model of the facility was fabricated. This 1/10th-scale model was used to gain further understanding of the flow characteristics in the IRT. The model was outfitted with instrumentation and data acquisition systems to determine pressures, velocities, and flow angles in the settling chamber and test section. Parametric flow quality studies involving the insertion and removal of a model of the IRT's distinctive heat exchanger (cooler) and/or of a honeycomb in the settling chamber were performed. These experiments illustrate the resulting improvement or degradation in flow quality.
NASA Technical Reports Server (NTRS)
Murch, Austin M.; Foster, John V.
2007-01-01
A simulation study was conducted to investigate aerodynamic modeling methods for prediction of post-stall flight dynamics of large transport airplanes. The research approach involved integrating dynamic wind tunnel data from rotary balance and forced oscillation testing with static wind tunnel data to predict aerodynamic forces and moments during highly dynamic departure and spin motions. Several state-of-the-art aerodynamic modeling methods were evaluated and predicted flight dynamics using these various approaches were compared. Results showed the different modeling methods had varying effects on the predicted flight dynamics and the differences were most significant during uncoordinated maneuvers. Preliminary wind tunnel validation data indicated the potential of the various methods for predicting steady spin motions.
Interference-free measurements of the subsonic aerodynamics of slanted-base ogive cylinders
NASA Technical Reports Server (NTRS)
Britcher, Colin P.; Alcorn, Charles W.
1991-01-01
Drag, lift, pitching moment, and base-pressure measurements have been made, free of support interference, on a range of slanted-base ogive cylinders, using the NASA Langley Research Center 13-in magnetic suspension and balance system. Test Mach numbers were in the range 0.04-0.2. Two types of wake flow were observed, a quasi-symmetric turbulent closure or a longitudinal vortex flow. Aerodynamic characteristics differ dramatically between the two wake types. Drag measurements are shown to be in agreement with previous tests. A hysteretic behavior of the wake with varying Reynold's number has been discovered for the 45-deg base. An interaction between forebody boundary-layer state and wake flow and base pressures has been detected for higher slant angles.
NASA Technical Reports Server (NTRS)
Brandon, Jay M.; Foster, John V.
1998-01-01
As airplane designs have trended toward the expansion of flight envelopes into the high angle of attack and high angular rate regimes, concerns regarding modeling the complex unsteady aerodynamics for simulation have arisen. Most current modeling methods still rely on traditional body axis damping coefficients that are measured using techniques which were intended for relatively benign flight conditions. This paper presents recent wind tunnel results obtained during large-amplitude pitch, roll and yaw testing of several fighter airplane configurations. A review of the similitude requirements for applying sub-scale test results to full-scale conditions is presented. Data is then shown to be a strong function of Strouhal number - both the traditional damping terms, but also the associated static stability terms. Additionally, large effects of sideslip are seen in the damping parameter that should be included in simulation math models. Finally, an example of the inclusion of frequency effects on the data in a simulation is shown.
Computation of rotor aerodynamic loads with a constant vorticity contour free wake model
NASA Technical Reports Server (NTRS)
Quackenbush, Todd R.; Wachspress, Daniel A.; Boschitsch, Alexander H.
1991-01-01
An analytical method is presented which facilitates the study of isolated rotors with an improved approach to wake simulation. Vortex filaments are simulated along contours of constant sheet strength for the sheet of vorticity resulting from each rotor blade. Curved vortex elements comprise the filaments which can be distorted by the local velocity field. Called the Constant Vorticity Contour wake model, the approach permits the simulation of the blades' wakes corresponding to the full span of the rotor blade. The discretization of the wake of the rotor blade produces spacing and structure that are consistent with the spatial and temporal variations in the loading. A vortex-lattice aerodynamic model of the blade is also included which introduces a finite-element structural model of the blade and consideration of the force and moment trim analysis. Results of the present version of the simulation, called RotorCRAFT, are found to correlate well with H-34 flight-test data.
NASA Technical Reports Server (NTRS)
Levine, J. J.
1999-01-01
This paper presents the terms of an Educational grant for Model Building 101. The terms of the grant includes the following: 1) 4 Training sessions of one week each (5 days/6 nights) at: Dryden, Langley, Lewis, and the California Museum of Science and Industry; 2) The sessions were to be attended by local educators, solicited and secured by NASA; 3) The cooperative program of MB101 and NASA was to set up a course for middle school students to learn aerodynamics through the building and flying of specialized small model airplanes. This program was already operating successfully on a local level through MB101 in Marietta, Georgia and was published monthly in Model Builder Magazine. MB101 supplies information for schools and groups throughout the country; and 4) Video and art department facilities of NASA were promised to be made available to MB101 for the preparation of instructional videos and preparation of training manuals.
Scaling of Lift Degradation Due to Anti-Icing Fluids Based Upon the Aerodynamic Acceptance Test
NASA Technical Reports Server (NTRS)
Broeren, Andy P.; Riley, James T.
2012-01-01
In recent years, the FAA has worked with Transport Canada, National Research Council Canada (NRC) and APS Aviation, Inc. to develop allowance times for aircraft operations in ice-pellet precipitation. These allowance times are critical to ensure safety and efficient operation of commercial and cargo flights. Wind-tunnel testing with uncontaminated anti-icing fluids and fluids contaminated with simulated ice pellets had been carried out at the NRC Propulsion and Icing Wind Tunnel (PIWT) to better understand the flowoff characteristics and resulting aerodynamic effects. The percent lift loss on the thin, high-performance wing model tested in the PIWT was determined at 8 angle of attack and used as one of the evaluation criteria in determining the allowance times. Because it was unclear as to how performance degradations measured on this model were relevant to an actual airplane configuration, some means of interpreting the wing model lift loss was deemed necessary. This paper describes how the lift loss was related to the loss in maximum lift of a Boeing 737-200ADV airplane through the Aerodynamic Acceptance Test (AAT) performed for fluids qualification. A loss in maximum lift coefficient of 5.24 percent on the B737-200ADV airplane (which was adopted as the threshold in the AAT) corresponds to a lift loss of 7.3 percent on the PIWT model at 8 angle of attack. There is significant scatter in the data used to develop the correlation related to varying effects of the anti-icing fluids that were tested and other factors. A statistical analysis indicated the upper limit of lift loss on the PIWT model was 9.2 percent. Therefore, for cases resulting in PIWT model lift loss from 7.3 to 9.2 percent, extra scrutiny of the visual observations is required in evaluating fluid performance with contamination.
Scaling of Lift Degradation Due to Anti-Icing Fluids Based Upon the Aerodynamic Acceptance Test
NASA Technical Reports Server (NTRS)
Broeren, Andy; Riley, Jim
2012-01-01
In recent years, the FAA has worked with Transport Canada, National Research Council Canada (NRC) and APS Aviation, Inc. to develop allowance times for aircraft operations in ice-pellet precipitation. These allowance times are critical to ensure safety and efficient operation of commercial and cargo flights. Wind-tunnel testing with uncontaminated anti-icing fluids and fluids contaminated with simulated ice pellets had been carried out at the NRC Propulsion and Icing Wind Tunnel (PIWT) to better understand the flow-off characteristics and resulting aerodynamic effects. The percent lift loss on the thin, high-performance wing model tested in the PIWT was determined at 8 deg. angle of attack and used as one of the evaluation criteria in determining the allowance times. Because it was unclear as to how performance degradations measured on this model were relevant to an actual airplane configuration, some means of interpreting the wing model lift loss was deemed necessary. This paper describes how the lift loss was related to the loss in maximum lift of a Boeing 737-200ADV airplane through the Aerodynamic Acceptance Test (AAT) performed for fluids qualification. A loss in maximum lift coefficient of 5.24% on the B737-200ADV airplane (which was adopted as the threshold in the AAT) corresponds to a lift loss of 7.3% on the PIWT model at 8 deg. angle of attack. There is significant scatter in the data used to develop the correlation related to varying effects of the anti-icing fluids that were tested and other factors. A statistical analysis indicated the upper limit of lift loss on the PIWT model was 9.2%. Therefore, for cases resulting in PIWT model lift loss from 7.3% to 9.2%, extra scrutiny of the visual observations is required in evaluating fluid performance with contamination.
1998-09-01
The CMARC panel-code is evaluated for the development of an aerodynamic model of the Naval Postgraduate School FROG Unmanned Air Vehicle (UAV). CMARC...model of the NPS FROG UAV is developed to obtain stability derivative data at the cruise flight condition. Emphasis is placed on comparing the CMARC data
A stochastic aerodynamic model for stationary blades in unsteady 3D wind fields
NASA Astrophysics Data System (ADS)
Fluck, Manuel; Crawford, Curran
2016-09-01
Dynamic loads play an important roll in the design of wind turbines, but establishing the life-time aerodynamic loads (e.g. extreme and fatigue loads) is a computationally expensive task. Conventional (deterministic) methods to analyze long term loads, which rely on the repeated analysis of multiple different wind samples, are usually too expensive to be included in optimization routines. We present a new stochastic approach, which solves the aerodynamic system equations (Lagrangian vortex model) in the stochastic space, and thus arrive directly at a stochastic description of the coupled loads along a turbine blade. This new approach removes the requirement of analyzing multiple different realizations. Instead, long term loads can be extracted from a single stochastic solution, a procedure that is obviously significantly faster. Despite the reduced analysis time, results obtained from the stochastic approach match deterministic result well for a simple test-case (a stationary blade). In future work, the stochastic method will be extended to rotating blades, thus opening up new avenues to include long term loads into turbine optimization.
A method of infrared imaging missile's aerodynamic heating modeling and simulations
NASA Astrophysics Data System (ADS)
Cao, Chunqin; Xiang, Jingbo; Zhang, Xiaoyang; Wang, Weiqiang
2013-09-01
The infrared (IR) imaging missile's dome will be heated when fly at high speed in the atmosphere because of the friction of the air flow blocking. The detector's performance will be decline if the dome surface is heated to a certain temperature. In this paper, we find a right way to evaluate the aerothermal effects in the imaging and information processing algorithm. Which have three steps including the aerothermal radiation calculation, quantization and image reconstruction. Firstly, the aerothermal radiation is calculated by using a combination of both methods of theoretical analysis and experiment data. Secondly, the relationship between aerothermal radiation and IR images background mean gray and noise can be calculated through the analysis of the experiment data. At last, we can rebuild an aerodynamic heating effect of infrared images fusion with target and decoy, which can be used for virtual prototyping platform missile trajectory simulation. It can be found that the above constructed images have good agreements with the actual image according to comparison between the simulation data and experiment data. It is an economic method that can solve the lab aerodynamic heating simulation and modeling problems.
NASA Astrophysics Data System (ADS)
Namba, Masanobu; Nishino, Ryohei
The purpose of this paper is to study the effect of neighboring blade rows on the unsteady aerodynamic response of oscillating cascade blades on the basis of a genuine three-dimensional model. To this end, mathematical formulations based on the lifting surface theory are developed for a pair of contra-rotating annular cascades of oscillating blades. The mechanism of frequency scattering of blade loadings and mode scattering of acoustic waves resulting from interaction between the blade rows in relative rotational motions is mathematically explained. Simultaneous integral equations for all frequency components of blade loadings are derived from the flow tangency condition on the blade surfaces of both blade rows. The validity of the computation codes is verified.
NASA Astrophysics Data System (ADS)
Poirier, Vincent
Mesh deformation schemes play an important role in numerical aerodynamic optimization. As the aerodynamic shape changes, the computational mesh must adapt to conform to the deformed geometry. In this work, an extension to an existing fast and robust Radial Basis Function (RBF) mesh movement scheme is presented. Using a reduced set of surface points to define the mesh deformation increases the efficiency of the RBF method; however, at the cost of introducing errors into the parameterization by not recovering the exact displacement of all surface points. A secondary mesh movement is implemented, within an adjoint-based optimization framework, to eliminate these errors. The proposed scheme is tested within a 3D Euler flow by reducing the pressure drag while maintaining lift of a wing-body configured Boeing-747 and an Onera-M6 wing. As well, an inverse pressure design is executed on the Onera-M6 wing and an inverse span loading case is presented for a wing-body configured DLR-F6 aircraft.
Aerodynamic characteristics of a powered tilt-proprotor wind tunnel model
NASA Technical Reports Server (NTRS)
Wilson, J. C.; Mineck, R. E.; Freeman, C. E.
1976-01-01
An investigation was conducted in the Langley V/STOL tunnel to determine the performance, stability and control, and rotor-wake interaction effects of a powered tilt-proprotor aircraft model with gimbal-hub rotors. Tests were conducted at representative flight conditions for hover, helicopter, transition, and airplane flight. Force and moment data were obtained for the complete model and for each of the two rotors. In addition to wind-speed variation, the angle of attack, angle of sideslip, rotor speed, rotor collective pitch, longitudinal cyclic pitch, rotor pylon angle, and configuration geometry were varied. The results, presented in graphical form, are available in tabular form to facilitate the validation of analytical methods of defining the aerodynamic characteristics of tilt-proprotor configurations.
A quasi-steady aerodynamic model for flapping flight with improved adaptability.
Lee, Y J; Lua, K B; Lim, T T; Yeo, K S
2016-04-28
An improved quasi-steady aerodynamic model for flapping wings in hover has been developed. The purpose of this model is to yield rapid predictions of lift generation and efficiency during the design phase of flapping wing micro air vehicles. While most existing models are tailored for a specific flow condition, the present model is applicable over a wider range of Reynolds number and Rossby number. The effects of wing aspect ratio and taper ratio are also considered. The model was validated by comparing against numerical simulations and experimental measurements. Wings with different geometries undergoing distinct kinematics at varying flow conditions were tested during validation. Generally, model predictions of mean force coefficients were within 10% of numerical simulation results, while the deviations in power coefficients could be up to 15%. The deviation is partly due to the model not taking into consideration the initial shedding of the leading-edge vortex and wing-wake interaction which are difficult to account under quasi-steady assumption. The accuracy of this model is comparable to other models in literature, which had to be specifically designed or tuned to a narrow range of operation. In contrast, the present model has the advantage of being applicable over a wider range of flow conditions without prior tuning or calibration, which makes it a useful tool for preliminary performance evaluations.
NASA Technical Reports Server (NTRS)
Jacobs, P. F.; Flechner, S. G.
1976-01-01
A baseline wing and a version of the same wing fitted with winglets were tested. The longitudinal aerodynamic characteristics were determined through an angle-of-attack range from -1 deg to 10 deg at an angle of sideslip of 0 deg for Mach numbers of 0.750, 0.800, and 0.825. The lateral aerodynamic characteristics were determined through the same angle-of-attack range at fixed sideslip angles of 2.5 deg and 5 deg. Both configurations were investigated at Reynolds numbers of 13,000,000, per meter (4,000,000 per foot) and approximately 20,000,000 per meter (6,000,000 per foot). The winglet configuration showed slight increases over the baseline wing in static longitudinal and lateral aerodynamic stability throughout the test Mach number range for a model design lift coefficient of 0.53. Reynolds number variation had very little effect on stability.
NASA Astrophysics Data System (ADS)
Shimada, Kenji; Ishihara, Takeshi
2012-01-01
It is well known that a bluff body cross-section exhibits various kinds of aerodynamic instabilities such as vortex-induced vibration, galloping and torsional flutter. Since these cross-sections are used in long-span bridges and tall buildings, it is important to predict their occurrence in wind resistant structural design. In this paper, the authors make a series of comparisons of unsteady wind forces, unsteady pressure distributions and free vibration responses between previously conducted studies and an unsteady two-dimensional k-ɛ model for rectangular cross-sections with cross-sectional ratios of 2 and 4 in a smooth uniform flow in order to verify computational predictability of aerodynamic instabilities. As a result, the computation successfully predicted the onset velocities and responses of these aerodynamic instabilities for these cross-sectional ratios, which are common to tall buildings and long bridges.
X based interactive computer graphics applications for aerodynamic design and education
NASA Technical Reports Server (NTRS)
Benson, Thomas J.; Higgs, C. Fred, III
1995-01-01
Six computer applications packages have been developed to solve a variety of aerodynamic problems in an interactive environment on a single workstation. The packages perform classical one dimensional analysis under the control of a graphical user interface and can be used for preliminary design or educational purposes. The programs were originally developed on a Silicon Graphics workstation and used the GL version of the FORMS library as the graphical user interface. These programs have recently been converted to the XFORMS library of X based graphics widgets and have been tested on SGI, IBM, Sun, HP and PC-Lunix computers. The paper will show results from the new VU-DUCT program as a prime example. VU-DUCT has been developed as an educational package for the study of subsonic open and closed loop wind tunnels.
NASA Technical Reports Server (NTRS)
Ciffone, D. L.; Robinson, G. H.
1973-01-01
An analysis of the influence of engine response characteristics on the approach and landing of an externally blown flap aircraft was conducted using flight simulator facilities. The configuration of the aerodynamic model is described. The aerodynamic characteristics as a function of angle of attack, thrust coefficient, and flap deflection are presented in tabular form and as graphs.
NASA Astrophysics Data System (ADS)
Gaunaa, Mac; Heinz, Joachim; Skrzypiński, Witold
2016-09-01
The crossflow principle is one of the key elements used in engineering models for prediction of the aerodynamic loads on wind turbine blades in standstill or blade installation situations, where the flow direction relative to the wind turbine blade has a component in the direction of the blade span direction. In the present work, the performance of the crossflow principle is assessed on the DTU 10MW reference blade using extensive 3D CFD calculations. Analysis of the computational results shows that there is only a relatively narrow region in which the crossflow principle describes the aerodynamic loading well. In some conditions the deviation of the predicted loadings can be quite significant, having a large influence on for instance the integral aerodynamic moments around the blade centre of mass; which is very important for single blade installation applications. The main features of these deviations, however, have a systematic behaviour on all force components, which in this paper is employed to formulate the first version of an engineering correction method to the crossflow principle applicable for wind turbine blades. The new correction model improves the agreement with CFD results for the key aerodynamic loads in crossflow situations. The general validity of this model for other blade shapes should be investigated in subsequent works.
NASA Technical Reports Server (NTRS)
Mcgrath, Brian E.; Neuhart, Dan H.; Gatlin, Gregory M.; Oneil, Pat
1994-01-01
A flat-plate wind tunnel model of an advanced fighter configuration was tested in the NASA LaRC Subsonic Basic Research Tunnel and the 16- by 24-inch Water Tunnel. The test objectives were to obtain and evaluate the low-speed longitudinal aerodynamic characteristics of a candidate configuration for the integration of several new innovative wing designs. The flat plate test allowed for the initial evaluation of the candidate planform and was designated as the baseline planform for the innovative wing design study. Low-speed longitudinal aerodynamic data were obtained over a range of freestream dynamic pressures from 7.5 psf to 30 psf (M = 0.07 to M = 0.14) and angles-of-attack from 0 to 40 deg. The aerodynamic data are presented in coefficient form for the lift, induced drag, and pitching moment. Flow-visualization results obtained were photographs of the flow pattern over the flat plate model in the water tunnel for angles-of-attack from 10 to 40 deg. The force and moment coefficients and the flow-visualization photographs showed the linear and nonlinear aerodynamic characteristics due to attached flow and vortical flow over the flat plate model. Comparison between experiment and linear theory showed good agreement for the lift and induced drag; however, the agreement was poor for the pitching moment.
Evangelista, Dennis; Cardona, Griselda; Guenther-Gleason, Eric; Huynh, Tony; Kwong, Austin; Marks, Dylan; Ray, Neil; Tisbe, Adrian; Tse, Kyle; Koehl, Mimi
2014-01-01
We report the effects of posture and morphology on the static aerodynamic stability and control effectiveness of physical models based on the feathered dinosaur, [Formula: see text]Microraptor gui, from the Cretaceous of China. Postures had similar lift and drag coefficients and were broadly similar when simplified metrics of gliding were considered, but they exhibited different stability characteristics depending on the position of the legs and the presence of feathers on the legs and the tail. Both stability and the function of appendages in generating maneuvering forces and torques changed as the glide angle or angle of attack were changed. These are significant because they represent an aerial environment that may have shifted during the evolution of directed aerial descent and other aerial behaviors. Certain movements were particularly effective (symmetric movements of the wings and tail in pitch, asymmetric wing movements, some tail movements). Other appendages altered their function from creating yaws at high angle of attack to rolls at low angle of attack, or reversed their function entirely. While [Formula: see text]M. gui lived after [Formula: see text]Archaeopteryx and likely represents a side experiment with feathered morphology, the general patterns of stability and control effectiveness suggested from the manipulations of forelimb, hindlimb and tail morphology here may help understand the evolution of flight control aerodynamics in vertebrates. Though these results rest on a single specimen, as further fossils with different morphologies are tested, the findings here could be applied in a phylogenetic context to reveal biomechanical constraints on extinct flyers arising from the need to maneuver.
Unsteady Aerodynamic Force Sensing from Measured Strain
NASA Technical Reports Server (NTRS)
Pak, Chan-Gi
2016-01-01
A simple approach for computing unsteady aerodynamic forces from simulated measured strain data is proposed in this study. First, the deflection and slope of the structure are computed from the unsteady strain using the two-step approach. Velocities and accelerations of the structure are computed using the autoregressive moving average model, on-line parameter estimator, low-pass filter, and a least-squares curve fitting method together with analytical derivatives with respect to time. Finally, aerodynamic forces over the wing are computed using modal aerodynamic influence coefficient matrices, a rational function approximation, and a time-marching algorithm. A cantilevered rectangular wing built and tested at the NASA Langley Research Center (Hampton, Virginia, USA) in 1959 is used to validate the simple approach. Unsteady aerodynamic forces as well as wing deflections, velocities, accelerations, and strains are computed using the CFL3D computational fluid dynamics (CFD) code and an MSC/NASTRAN code (MSC Software Corporation, Newport Beach, California, USA), and these CFL3D-based results are assumed as measured quantities. Based on the measured strains, wing deflections, velocities, accelerations, and aerodynamic forces are computed using the proposed approach. These computed deflections, velocities, accelerations, and unsteady aerodynamic forces are compared with the CFL3D/NASTRAN-based results. In general, computed aerodynamic forces based on the lifting surface theory in subsonic speeds are in good agreement with the target aerodynamic forces generated using CFL3D code with the Euler equation. Excellent aeroelastic responses are obtained even with unsteady strain data under the signal to noise ratio of -9.8dB. The deflections, velocities, and accelerations at each sensor location are independent of structural and aerodynamic models. Therefore, the distributed strain data together with the current proposed approaches can be used as distributed deflection
Determining aerodynamic coefficients from high speed video of a free-flying model in a shock tunnel
NASA Astrophysics Data System (ADS)
Neely, Andrew J.; West, Ivan; Hruschka, Robert; Park, Gisu; Mudford, Neil R.
2008-11-01
This paper describes the application of the free flight technique to determine the aerodynamic coefficients of a model for the flow conditions produced in a shock tunnel. Sting-based force measurement techniques either lack the required temporal response or are restricted to large complex models. Additionally the free flight technique removes the flow interference produced by the sting that is present for these other techniques. Shock tunnel test flows present two major challenges to the practical implementation of the free flight technique. These are the millisecond-order duration of the test flows and the spatial and temporal nonuniformity of these flows. These challenges are overcome by the combination of an ultra-high speed digital video camera to record the trajectory, with spatial and temporal mapping of the test flow conditions. Use of a lightweight model ensures sufficient motion during the test time. The technique is demonstrated using the simple case of drag measurement on a spherical model, free flown in a Mach 10 shock tunnel condition.
NASA Astrophysics Data System (ADS)
Zhu, Xiaowei; Iungo, G. Valerio; Leonardi, Stefano; Anderson, William
2017-02-01
For a horizontally homogeneous, neutrally stratified atmospheric boundary layer (ABL), aerodynamic roughness length, z_0, is the effective elevation at which the streamwise component of mean velocity is zero. A priori prediction of z_0 based on topographic attributes remains an open line of inquiry in planetary boundary-layer research. Urban topographies - the topic of this study - exhibit spatial heterogeneities associated with variability of building height, width, and proximity with adjacent buildings; such variability renders a priori, prognostic z_0 models appealing. Here, large-eddy simulation (LES) has been used in an extensive parametric study to characterize the ABL response (and z_0) to a range of synthetic, urban-like topographies wherein statistical moments of the topography have been systematically varied. Using LES results, we determined the hierarchical influence of topographic moments relevant to setting z_0. We demonstrate that standard deviation and skewness are important, while kurtosis is negligible. This finding is reconciled with a model recently proposed by Flack and Schultz (J Fluids Eng 132:041203-1-041203-10, 2010), who demonstrate that z_0 can be modelled with standard deviation and skewness, and two empirical coefficients (one for each moment). We find that the empirical coefficient related to skewness is not constant, but exhibits a dependence on standard deviation over certain ranges. For idealized, quasi-uniform cubic topographies and for complex, fully random urban-like topographies, we demonstrate strong performance of the generalized Flack and Schultz model against contemporary roughness correlations.
NASA Technical Reports Server (NTRS)
Spekreijse, S. P.; Boerstoel, J. W.; Vitagliano, P. L.; Kuyvenhoven, J. L.
1992-01-01
About five years ago, a joint development was started of a flow simulation system for engine-airframe integration studies on propeller as well as jet aircraft. The initial system was based on the Euler equations and made operational for industrial aerodynamic design work. The system consists of three major components: a domain modeller, for the graphical interactive subdivision of flow domains into an unstructured collection of blocks; a grid generator, for the graphical interactive computation of structured grids in blocks; and a flow solver, for the computation of flows on multi-block grids. The industrial partners of the collaboration and NLR have demonstrated that the domain modeller, grid generator and flow solver can be applied to simulate Euler flows around complete aircraft, including propulsion system simulation. Extension to Navier-Stokes flows is in progress. Delft Hydraulics has shown that both the domain modeller and grid generator can also be applied successfully for hydrodynamic configurations. An overview is given about the main aspects of both domain modelling and grid generation.
A flight experiment to measure rarefied-flow aerodynamics
NASA Technical Reports Server (NTRS)
Blanchard, Robert C.
1990-01-01
A flight experiment to measure rarefied-flow aerodynamics of a blunt lifting body is being developed by NASA. This experiment, called the Rarefied-Flow Aerodynamic Measurement Experiment (RAME), is part of the Aeroassist Flight Experiment (AFE) mission, which is a Pathfinder design tool for aeroassisted orbital transfer vehicles. The RAME will use flight measurements from accelerometers, rate gyros, and pressure transducers, combined with knowledge of AFE in-flight mass properties and trajectory, to infer aerodynamic forces and moments in the rarefied-flow environment, including transition into the hypersonic continuum regime. Preflight estimates of the aerodynamic measurements are based upon environment models, existing computer simulations, and ground test results. Planned maneuvers at several altitudes will provide a first-time opportunity to examine gas-surface accommondation effects on aerodynamic coefficients in an environment of changing atmospheric composition. A description is given of the RAME equipment design.
NASA Technical Reports Server (NTRS)
Stoliker, Patrick C.; Bosworth, John T.; Georgie, Jennifer
1997-01-01
The X-31A aircraft has a unique configuration that uses thrust-vector vanes and aerodynamic control effectors to provide an operating envelope to a maximum 70 deg angle of attack, an inherently nonlinear portion of the flight envelope. This report presents linearized versions of the X-31A longitudinal and lateral-directional control systems, with aerodynamic models sufficient to evaluate characteristics in the poststall envelope at 30 deg, 45 deg, and 60 deg angle of attack. The models are presented with detail sufficient to allow the reader to reproduce the linear results or perform independent control studies. Comparisons between the responses of the linear models and flight data are presented in the time and frequency domains to demonstrate the strengths and weaknesses of the ability to predict high-angle-of-attack flight dynamics using linear models. The X-31A six-degree-of-freedom simulation contains a program that calculates linear perturbation models throughout the X-31A flight envelope. The models include aerodynamics and flight control system dynamics that are used for stability, controllability, and handling qualities analysis. The models presented in this report demonstrate the ability to provide reasonable linear representations in the poststall flight regime.
NASA Astrophysics Data System (ADS)
Rinehart, Taylor Jay
Wind turbine sizes have been steadily increasing to reduce the cost of generating electricity using wind energy. The increased wind turbine blade size has led to increased interest in the accurate prediction of the aerodynamics of large wind turbine blades. In this work, two-dimensional simulations of wind turbine airfoils and three-dimensional simulations of the Sandia 100 m wind turbine blade were conducted. The focus of the simulations was to evaluate improvements in turbulence modeling for wind turbine applications. The flow field was modeled using a Reynolds-Averaged Navier-Stokes flow solver. The turbulence model included transition modeling to capture the significant regions of laminar flow found on wind turbine airfoils and wind turbine blades. The turbulence model was also modified to increase sensitivity to adverse pressure gradients. The effects of modifying the turbulence modeling were quantified using lift and drag for two-dimensional simulations while wind turbine thrust and power were used as metrics for three-dimensional simulations. The two-dimensional studies showed that the adverse pressure gradient correction lowered lift predictions post-stall by about 13%, significantly reducing lift over-prediction and bringing simulations closer to experimental results. Transition modeling lowered drag predictions by 30% to 50% at low angles of attack bringing the predicted values into good agreement with experimental results. The addition of transition modeling in the three-dimensional simulations increased the predicted thrust by 1% to 3% and predicted power by 3% to 6%. The extent of laminar flow was visualized using intermittency. Laminar flow was observed on large portions of the Sandia 100 m blade at normal operating conditions. A preliminary study on the effects of leading edge tubercles on the Sandia 100 m blade was performed, no significant changes in wind turbine performance were observed at nominal operating conditions.
Han, Jong-Seob; Kim, Joong-Kwan; Chang, Jo Won; Han, Jae-Hung
2015-07-30
A quasi-steady aerodynamic model in consideration of the center of pressure (C.P.) was developed for insect flight. A dynamically scaled-up robotic hawkmoth wing was used to obtain the translational lift, drag, moment and rotational force coefficients. The translational force coefficients were curve-fitted with respect to the angles of attack such that two coefficients in the Polhamus leading-edge suction analogy model were obtained. The rotational force coefficient was also compared to that derived by the standard Kutta-Joukowski theory. In order to build the accurate pitching moment model, the locations of the C.Ps. and its movements depending on the pitching velocity were investigated in detail. We found that the aerodynamic moment model became suitable when the rotational force component was assumed to act on the half-chord. This implies that the approximation borrowed from the conventional airfoil concept, i.e., the 'C.P. at the quarter-chord' may lead to an incorrect moment prediction. In the validation process, the model showed excellent time-course force and moment estimations in comparison with the robotic wing measurement results. A fully nonlinear multibody flight dynamic simulation was conducted to check the effect of the traveling C.P. on the overall flight dynamics. This clearly showed the importance of an accurate aerodynamic moment model.
Xiao, Haosu; Zuo, Baojun; Tian, Yi; Zhang, Wang; Hao, Chenglong; Liu, Chaofeng; Li, Qi; Li, Fan; Zhang, Li; Fan, Zhigang
2012-12-20
We investigated the joint influences exerted by the nonuniform aerodynamic flow field surrounding the optical dome and the aerodynamic heating of the dome on imaging quality degradation of an airborne optical system. The Spalart-Allmaras model provided by FLUENT was used for flow computations. The fourth-order Runge-Kutta algorithm based ray tracing program was used to simulate optical transmission through the aerodynamic flow field and the dome. Four kinds of imaging quality evaluation parameters were presented: wave aberration of the exit pupil, point spread function, encircled energy, and modulation transfer function. The results show that the aero-optical disturbance of the aerodynamic flow field and the aerodynamic heating of the dome significantly affect the imaging quality of an airborne optical system.
NASA Technical Reports Server (NTRS)
Nelson, D. P.
1981-01-01
Tabulated aerodynamic data from coannular nozzle performance tests are given for test runs 26 through 37. The data include nozzle thrust coefficient parameters, nozzle discharge coefficients, and static pressure tap measurements.
Latest results from the EU project AVATAR: Aerodynamic modelling of 10 MW wind turbines
NASA Astrophysics Data System (ADS)
Ceyhan, J. G. Schepers O.; Boorsma, K.; Gonzalez, A.; Munduate, X.; Pires, O.; Sørensen, N..; Ferreira, C.; Sieros, G.; Madsen, J.; Voutsinas, S.; Lutz, T.; Barakos, G.; Colonia, S.; Heißelmann, H.; Meng, F.; Croce, A.
2016-09-01
This paper presents the most recent results from the EU project AVATAR in which aerodynamic models are improved and validated for wind turbines on a scale of 10 MW and more. Measurements on a DU 00-W-212 airfoil are presented which have been taken in the pressurized DNW-HDG wind tunnel up to a Reynolds number of 15 Million. These measurements are compared with measurements in the LM wind tunnel for Reynolds numbers of 3 and 6 Million and with calculational results. In the analysis of results special attention is paid to high Reynolds numbers effects. CFD calculations on airfoil performance showed an unexpected large scatter which eventually was reduced by paying even more attention to grid independency and domain size in relation to grid topology. Moreover calculations are presented on flow devices (leading and trailing edge flaps and vortex generators). Finally results are shown between results from 3D rotor models where a comparison is made between results from vortex wake methods and BEM methods at yawed conditions.
The effect of plasma actuator on the depreciation of the aerodynamic drag on box model
NASA Astrophysics Data System (ADS)
Harinaldi, Budiarso, Julian, James; Rabbani M., N.
2016-06-01
Recent active control research advances have provided many benefits some of which in the field of transportation by land, sea as well as by air. Flow engineering by using active control has proven advantages in energy saving significantly. One of the active control equipment that is being developed, especially in the 21st century, is a plasma actuator, with the ability to modify the flow of fluid by the approach of ion particles makes these actuators a very powerful and promising tool. This actuator can be said to be better to the previously active control such as suction, blowing and synthetic jets because it is easier to control, more flexible because it has no moving parts, easy to be manufactured and installed, and consumes a small amount of energy with maximum capability. Plasma actuator itself is the composition of a material composed of copper and a dielectric sheet, where the copper sheets act as an electricity conductor and the dielectric sheet as electricity insulator. Products from the plasma actuators are ion wind which is the result of the suction of free air around the actuator to the plasma zone. This study investigates the ability of plasma actuators in lowering aerodynamic drag which is commonly formed in the models of vehicles by varying the shape of geometry models and the flow speed.
MacKinnon, D.J.; Clow, G.D.; Tigges, R.K.; Reynolds, R.L.; Chavez, P.S.
2004-01-01
The vulnerability of dryland surfaces to wind erosion depends importantly on the absence or the presence and character of surface roughness elements, such as plants, clasts, and topographic irregularities that diminish wind speed near the surface. A model for the friction velocity ratio has been developed to account for wind sheltering by many different types of co-existing roughness elements. Such conditions typify a monitored area in the central Mojave Desert, California, that experiences frequent sand movement and dust emission. Two additional models are used to convert the friction velocity ratio to the surface roughness length (zo) for momentum. To calculate roughness lengths from these models, measurements were made at 11 sites within the monitored area to characterize the surface roughness element. Measurements included (1) the number of roughness species (e.g., plants, small-scale topography, clasts), and their associated heights and widths, (2) spacing among species, and (3) vegetation porosity (a measurement of the spatial distribution of woody elements of a plant). Documented or estimated values of drag coefficients for different species were included in the modeling. At these sites, wind-speed profiles were measured during periods of neutral atmospheric stability using three 9-m towers with three or four calibrated anemometers on each. Modeled roughness lengths show a close correspondence (correlation coefficient, 0.84-0.86) to the aerodynamically determined values at the field sites. The geometric properties of the roughness elements in the model are amenable to measurement at much higher temporal and spatial resolutions using remote-sensing techniques than can be accomplished through laborious ground-based methods. A remote-sensing approach to acquire values of the modeled roughness length is particularly important for the development of linked surface/atmosphere wind-erosion models sensitive to climate variability and land-use changes in areas such
NASA Technical Reports Server (NTRS)
Swihert, John M
1958-01-01
A brief investigation of a target-type thrust reverser on a single-engine fighter model has been conducted in the Langley 16-foot transonic tunnel at Mach numbers from 0.20 to 1.05.At Mach numbers of 0.80, 0.92, and 1.05, a hydrogen peroxide turbojet-engine simulator was operated with the thrust reverser extended. The angle of attack was varied from 0 degrees to 5 degrees at these Mach numbers. The Reynolds number of the free stream, based on the mean aerodynamic chord, was about 5 x 10(6). It was estimated that reversed jet operations separated the model boundary-layer flow over the upper surface of the horizontal tail and upper part of the afterbody. This resulted in a positive pitch increment due to reversed jet operation. Jet-on operation also tended to stabilize the severe lateral oscillations which occurred with the reverser extended and the jet off. It appeared that these jet-off oscillations were the result of an alternating separation and reattachment of the flow on the rearmost portions of the fuselage afterbody.
Bailly, Lucie; Henrich, Nathalie; Pelorson, Xavier
2010-05-01
Occurrences of period-doubling are found in human phonation, in particular for pathological and some singing phonations such as Sardinian A Tenore Bassu vocal performance. The combined vibration of the vocal folds and the ventricular folds has been observed during the production of such low pitch bass-type sound. The present study aims to characterize the physiological correlates of this acoustical production and to provide a better understanding of the physical interaction between ventricular fold vibration and vocal fold self-sustained oscillation. The vibratory properties of the vocal folds and the ventricular folds during phonation produced by a professional singer are analyzed by means of acoustical and electroglottographic signals and by synchronized glottal images obtained by high-speed cinematography. The periodic variation in glottal cycle duration and the effect of ventricular fold closing on glottal closing time are demonstrated. Using the detected glottal and ventricular areas, the aerodynamic behavior of the laryngeal system is simulated using a simplified physical modeling previously validated in vitro using a larynx replica. An estimate of the ventricular aperture extracted from the in vivo data allows a theoretical prediction of the glottal aperture. The in vivo measurements of the glottal aperture are then compared to the simulated estimations.
Aerodynamic characteristics of a 1/6-scale powered model of the rotor systems research aircraft
NASA Technical Reports Server (NTRS)
Mineck, R. E.; Freeman, C. E.
1977-01-01
A wind-tunnel investigation was conducted to determine the effects of the main-rotor wake on the aerodynamic characteristics of the rotor systems research aircraft (RSRA). For the investigation, a 1/6-scale model with a four-blade articulated main rotor was used. Tests were conducted with and without the main rotor. Both the helicopter and the compound helicopter were tested. The latter configuration included the auxiliary thrust engines and the variable-incidence wing. Data were obtained over ranges of angle of attack, angle of sideslip, and main-rotor collective pitch angle at several main-rotor advance ratios. Results are presented for the total loads on the airframe as well as the loads on the rotor, the wing, and the tail. The results indicated that without the effect of the rotor wake, the RSRA had static longitudinal and directional stability and positive effective dihedral. With the effect of the main rotor and its wake, the RSRA exhibited longitudinal instability but retained static directional stability and positive effective dihedral.
Acoustic and aerodynamic study of a pusher-propeller aircraft model
NASA Technical Reports Server (NTRS)
Soderman, Paul T.; Horne, W. Clifton
1990-01-01
An aerodynamic and acoustic study was made of a pusher-propeller aircraft model in the NASA-Ames 7 x 10 ft Wind Tunnel. The test section was changed to operate as an open jet. The 591 mm diameter unswept propeller was operated alone and in the wake of three empennages: an I tail, Y tail, and a V tail. The radiated noise and detailed wake properties were measured. Results indicate that the unsteady blade loading caused by the blade interactions with the wake mean velocity distribution had a strong effect on the harmonics of blade passage noise. The blade passage harmonics above the first were substantially increased in all horizontal directions by the empennage/propeller interaction. Directivity in the plane of the propeller was maximum perpendicular to the blade surface. Increasing the tail loading caused the propeller harmonics to increase 3 to 5 dB for an empennage/propeller spacing of 0.38 mean empennage chords. The interaction noise became weak as empennage propeller spacing was increased beyond 1.0 mean empennage chord lengths. Unlike the mean wake deficit, the wake turbulence had only a small effect on the propeller noise, that effect being a small increase in the broadband noise.
Acoustic and aerodynamic testing of a scale model variable pitch fan
NASA Technical Reports Server (NTRS)
Jutras, R. R.; Kazin, S. B.
1974-01-01
A fully reversible pitch scale model fan with variable pitch rotor blades was tested to determine its aerodynamic and acoustic characteristics. The single-stage fan has a design tip speed of 1160 ft/sec (353.568 m/sec) at a bypass pressure ratio of 1.5. Three operating lines were investigated. Test results show that the blade pitch for minimum noise also resulted in the highest efficiency for all three operating lines at all thrust levels. The minimum perceived noise on a 200-ft (60.96 m) sideline was obtained with the nominal nozzle. At 44% of takeoff thrust, the PNL reduction between blade pitch and minimum noise blade pitch is 1.8 PNdB for the nominal nozzle and decreases with increasing thrust. The small nozzle (6% undersized) has the highest efficiency at all part thrust conditions for the minimum noise blade pitch setting; although, the noise is about 1.0 PNdB higher for the small nozzle at the minimum noise blade pitch position.
NASA Technical Reports Server (NTRS)
Lummus, J. R.; Joyce, G. T.; Omalley, C. D.
1980-01-01
The ability of current methodologies to accurately predict the aerodynamic characteristics identified as uncertainties was evaluated for two aircraft configurations. The two wind tunnel models studied horizontal altitude takeoff and landing V/STOL fighter aircraft derivatives.
Defraeye, Thijs; Blocken, Bert; Koninckx, Erwin; Hespel, Peter; Carmeliet, Jan
2010-08-26
This study aims at assessing the accuracy of computational fluid dynamics (CFD) for applications in sports aerodynamics, for example for drag predictions of swimmers, cyclists or skiers, by evaluating the applied numerical modelling techniques by means of detailed validation experiments. In this study, a wind-tunnel experiment on a scale model of a cyclist (scale 1:2) is presented. Apart from three-component forces and moments, also high-resolution surface pressure measurements on the scale model's surface, i.e. at 115 locations, are performed to provide detailed information on the flow field. These data are used to compare the performance of different turbulence-modelling techniques, such as steady Reynolds-averaged Navier-Stokes (RANS), with several k-epsilon and k-omega turbulence models, and unsteady large-eddy simulation (LES), and also boundary-layer modelling techniques, namely wall functions and low-Reynolds number modelling (LRNM). The commercial CFD code Fluent 6.3 is used for the simulations. The RANS shear-stress transport (SST) k-omega model shows the best overall performance, followed by the more computationally expensive LES. Furthermore, LRNM is clearly preferred over wall functions to model the boundary layer. This study showed that there are more accurate alternatives for evaluating flow around bluff bodies with CFD than the standard k-epsilon model combined with wall functions, which is often used in CFD studies in sports.
NASA Technical Reports Server (NTRS)
Queijo, M. J.; Wells, W. R.; Keskar, D. A.
1979-01-01
A simple vortex system, used to model unsteady aerodynamic effects into the rigid body longitudinal equations of motion of an aircraft, is described. The equations are used in the development of a parameter extraction algorithm. Use of the two parameter-estimation modes, one including and the other omitting unsteady aerodynamic modeling, is discussed as a means of estimating some acceleration derivatives. Computer generated data and flight data, used to demonstrate the use of the parameter-extraction algorithm are studied.
Effect of body aerodynamics on the dynamic flight stability of the hawkmoth Manduca sexta.
Nguyen, Anh Tuan; Han, Jong-Seob; Han, Jae-Hung
2016-12-14
This study explores the effects of the body aerodynamics on the dynamic flight stability of an insect at various different forward flight speeds. The insect model, whose morphological parameters are based on measurement data from the hawkmoth Manduca sexta, is treated as an open-loop six-degree-of-freedom dynamic system. The aerodynamic forces and moments acting on the insect are computed by an aerodynamic model that combines the unsteady panel method and the extended unsteady vortex-lattice method. The aerodynamic model is then coupled to a multi-body dynamic code to solve the system of motion equations. First, the trimmed flight conditions of insect models with and without consideration of the body aerodynamics are obtained using a trim search algorithm. Subsequently, the effects of the body aerodynamics on the dynamic flight stability are analysed through modal structures, i.e., eigenvalues and eigenvectors in this case, which are based on linearized equations of motion. The solutions from the nonlinear and linearized equations of motion due to gust disturbances are obtained, and the effects of the body aerodynamics are also investigated through these solutions. The results showed the important effect of the body aerodynamics at high-speed forward flight (in this paper at 4.0 and 5.0 m s(-1)) and the movement trends of eigenvalues when the body aerodynamics is included.
NASA Technical Reports Server (NTRS)
Aiken, T. N.
1973-01-01
An investigation was made of the static, wind-on aerodynamic and static noise characteristics of an augmentor wing having lobe type nozzles. The study was made in the Ames 7-by 10-Foot No. 1 Wind Tunnel using a small-scale, quasi-two-dimensional model. Several configurations of lobe nozzles as well as a normal slot nozzle were tested. Results indicate that lobe nozzles offer improved static and wind-on aerodynamics and reduced static noise relative to slot nozzles. Best wind-on performance was obtained when the tertiary gap was closed even though the static thrust augmentation was maximum with the gap open. Static thrust augmentation, wind-on lift and drag, and static noise directivity are presented as well as typical static and wind-on exit velocity profiles, surface pressure distributions and noise spectrums. The data are presented with limited discussion.
Van Dam, C P; Nakafuji, D Y; Bauer, C; Chao, D; Standish, K
2002-11-01
A computational design and analysis of a microtab based aerodynamic loads control system is presented. The microtab consists of a small tab that emerges from a wing approximately perpendicular to its surface in the vicinity of its trailing edge. Tab deployment on the upper side of the wing causes a decrease in the lift generation whereas deployment on the pressure side causes an increase. The computational methods applied in the development of this concept solve the governing Reynolds-averaged Navier-Stokes equations on structured, overset grids. The application of these methods to simulate the flows over lifting surface including the tabs has been paramount in the development of these devices. The numerical results demonstrate the effectiveness of the microtab and that it is possible to carry out a sensitivity analysis on the positioning and sizing of the tabs before they are implemented in successfully controlling the aerodynamic loads.
NASA Astrophysics Data System (ADS)
Raskin, Boris
Scaled wind tunnel models are necessary for the development of aircraft and spacecraft to simulate aerodynamic behavior. This allows for testing multiple iterations of a design before more expensive full-scale aircraft and spacecraft are built. However, the cost of building wind tunnel models can still be high because they normally require costly subtractive manufacturing processes, such as machining, which can be time consuming and laborious due to the complex surfaces of aerodynamic models. Rapid prototyping, commonly known as 3D printing, can be utilized to save on wind tunnel model manufacturing costs. A rapid prototype multi-material wind tunnel model was manufactured for this thesis to investigate the possibility of using PolyJet 3D printing to create a model that exhibits aeroelastic behavior. The model is of NASA's Adaptable Deployable entry and Placement (ADEPT) aerodynamic decelerator, used to decelerate a spacecraft during reentry into a planet's atmosphere. It is a 60° cone with a spherically blunted nose that consists of a 12 flexible panels supported by a rigid structure of nose, ribs, and rim. The novel rapid prototype multi-material model was instrumented and tested in two flow conditions. Quantitative comparisons were made of the average forces and dynamic forces on the model, demonstrating that the model matched expected behavior for average drag, but not Strouhal number, indicating that there was no aeroelastic behavior in this particular case. It was also noted that the dynamic properties (e.g., resonant frequency) associated with the mounting scheme are very important and may dominate the measured dynamic response.
The compressible aerodynamics of rotating blades based on an acoustic formulation
NASA Technical Reports Server (NTRS)
Long, L. N.
1983-01-01
An acoustic formula derived for the calculation of the noise of moving bodies is applied to aerodynamic problems. The acoustic formulation is a time domain result suitable for slender wings and bodies moving at subsonic speeds. A singular integral equation is derived in terms of the surface pressure which must then be solved numerically for aerodynamic purposes. However, as the 'observer' is moved onto the body surface, the divergent integrals in the acoustic formulation are semiconvergent. The procedure for regularization (or taking principal values of divergent integrals) is explained, and some numerical examples for ellipsoids, wings, and lifting rotors are presented. The numerical results show good agreement with available measured surface pressure data.
NASA Technical Reports Server (NTRS)
Wahls, R. A.; Adcock, J. B.; Witkowski, D. P.; Wright, F. L.
1995-01-01
A high Reynolds number investigation of a commercial transport model was conducted in the National Transonic Facility (NTF) at Langley Research Center. This investigation was part of a cooperative effort to test a 0.03-scale model of a Boeing 767 airplane in the NTF over a Mach number range of 0.70 to 0.86 and a Reynolds number range of 2.38 to 40.0 x 10(exp 6) based on the mean aerodynamic chord. One of several specific objectives of the current investigation was to evaluate the level of data repeatability attainable in the NTF. Data repeatability studies were performed at a Mach number of 0.80 with Reynolds numbers of 2.38, 4.45, and 40.0 x 10(exp 6) and also at a Mach number of 0.70 with a Reynolds number of 40.0 x 10(exp 6). Many test procedures and data corrections are addressed in this report, but the data presented do not include corrections for wall interference, model support interference, or model aeroelastic effects. Application of corrections for these three effects would not affect the results of this study because the corrections are systematic in nature and are more appropriately classified as sources of bias error. The repeatability of the longitudinal stability-axis force and moment data has been accessed. Coefficients of lift, drag, and pitching moment are shown to repeat well within the pretest goals of plus or minus 0.005, plus or minus 0.0001, and plus or minus 0.001, respectively, at a 95-percent confidence level over both short- and near-term periods.
NASA Technical Reports Server (NTRS)
Petynia, William W.; Croom, Delwin R.; Davenport, Edwin E.
1958-01-01
The low-speed aerodynamic and hydrodynamic characteristics of a proposed multijet water-based aircraft configuration for supersonic operation have been investigated. The design features include upward-rotating engines, body indentation, a single hydro-ski, and a wing with an aspect ratio of 3.0, a taper ratio of 0.143, 36.90 sweepback of the quarter-chord line, and NACA 65AO04 airfoil sections. For the aerodynamic investigation, with the flaps retracted, the model was longitudinally and directionally stable up to the stall. The all-movable horizontal tail was capable of trimming the model up to a lift coefficient of approximately 0.87. All flap configurations investigated had a tendency to become longitudinally unstable at stall. The effectiveness of the all-movable horizontal tail increased with increasing lift coefficient for all flap configurations investigated; however, with the large static margin of the configuration with the center of gravity at 0.25 mean aerodynamic chord, the all-movable horizontal tail was not powerful enough to trim all the various flapped configurations investigated throughout the angle-of-attack range. For the hydrodynamic investigation, longitudinal stability during take-offs and landings was satisfactory. Decreasing the area of the hydro-ski 60 percent increased the maximum resistance and emergence speed 40 and 70 percent, respectively. Without the jet exhaust, the resistance was reduced by simulating the vertical-lift component of the forward engines rotated upward. However, the jet exhaust of the forward engines increased the maximum resistance approximately 60 percent. The engine inlets and horizontal tail were free from spray for all loads investigated and for both hydro-ski sizes.
Aerodynamic performance of a hovering hawkmoth with flexible wings: a computational approach.
Nakata, Toshiyuki; Liu, Hao
2012-02-22
Insect wings are deformable structures that change shape passively and dynamically owing to inertial and aerodynamic forces during flight. It is still unclear how the three-dimensional and passive change of wing kinematics owing to inherent wing flexibility contributes to unsteady aerodynamics and energetics in insect flapping flight. Here, we perform a systematic fluid-structure interaction based analysis on the aerodynamic performance of a hovering hawkmoth, Manduca, with an integrated computational model of a hovering insect with rigid and flexible wings. Aerodynamic performance of flapping wings with passive deformation or prescribed deformation is evaluated in terms of aerodynamic force, power and efficiency. Our results reveal that wing flexibility can increase downwash in wake and hence aerodynamic force: first, a dynamic wing bending is observed, which delays the breakdown of leading edge vortex near the wing tip, responsible for augmenting the aerodynamic force-production; second, a combination of the dynamic change of wing bending and twist favourably modifies the wing kinematics in the distal area, which leads to the aerodynamic force enhancement immediately before stroke reversal. Moreover, an increase in hovering efficiency of the flexible wing is achieved as a result of the wing twist. An extensive study of wing stiffness effect on aerodynamic performance is further conducted through a tuning of Young's modulus and thickness, indicating that insect wing structures may be optimized not only in terms of aerodynamic performance but also dependent on many factors, such as the wing strength, the circulation capability of wing veins and the control of wing movements.
Unsteady Aerodynamic Testing Using the Dynamic Plunge Pitch and Roll Model Mount
NASA Technical Reports Server (NTRS)
Lutze, Frederick H.; Fan, Yigang
1999-01-01
A final report on the DyPPiR tests that were run are presented. Essentially it consists of two parts, a description of the data reduction techniques and the results. The data reduction techniques include three methods that were considered: 1) signal processing of wind on - wind off data; 2) using wind on data in conjunction with accelerometer measurements; and 3) using a dynamic model of the sting to predict the sting oscillations and determining the aerodynamic inputs using an optimization process. After trying all three, we ended up using method 1, mainly because of its simplicity and our confidence in its accuracy. The results section consists of time history plots of the input variables (angle of attack, roll angle, and/or plunge position) and the corresponding time histories of the output variables, C(sub L), C(sub D), C(sub m), C(sub l), C(sub m), C(sub n). Also included are some phase plots of one or more of the output variable vs. an input variable. Typically of interest are pitch moment coefficient vs. angle of attack for an oscillatory motion where the hysteresis loops can be observed. These plots are useful to determine the "more interesting" cases. Samples of the data as it appears on the disk are presented at the end of the report. The last maneuver, a rolling pull up, is indicative of the unique capabilities of the DyPPiR, allowing combinations of motions to be exercised at the same time.
Computational aerodynamics and design
NASA Technical Reports Server (NTRS)
Ballhaus, W. F., Jr.
1982-01-01
The role of computational aerodynamics in design is reviewed with attention given to the design process; the proper role of computations; the importance of calibration, interpretation, and verification; the usefulness of a given computational capability; and the marketing of new codes. Examples of computational aerodynamics in design are given with particular emphasis on the Highly Maneuverable Aircraft Technology. Finally, future prospects are noted, with consideration given to the role of advanced computers, advances in numerical solution techniques, turbulence models, complex geometries, and computational design procedures. Previously announced in STAR as N82-33348
Nonlinear aerodynamic wing design
NASA Technical Reports Server (NTRS)
Bonner, Ellwood
1985-01-01
The applicability of new nonlinear theoretical techniques is demonstrated for supersonic wing design. The new technology was utilized to define outboard panels for an existing advanced tactical fighter model. Mach 1.6 maneuver point design and multi-operating point compromise surfaces were developed and tested. High aerodynamic efficiency was achieved at the design conditions. A corollary result was that only modest supersonic penalties were incurred to meet multiple aerodynamic requirements. The nonlinear potential analysis of a practical configuration arrangement correlated well with experimental data.
NASA Technical Reports Server (NTRS)
Salters, L. B., Jr.; Schmeer, J. W.
1973-01-01
The aerodynamic and propulsion characteristics of a 1/6-scale propulsive-wing V/STOL air-powered model was investigated over the Mach number range from 0.40 to 0.96 and at angles of attack from -5 deg to 15 deg for several fan rotational speeds. Three fanduct-exit configurations were tested, including two exit areas. The model with 25-percent-thick wing had a drag-rise Mach number of 0.85, which is typical of aircraft with thinner, conventional, unswept wings.
NASA Technical Reports Server (NTRS)
Robinson, Ross B.; Morris, Odell A.
1960-01-01
An investigation has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel to determine the aerodynamic characteristics in pitch of a two-stage-rocket model configuration which simulated the last two stages of the launching vehicle for an inflatable sphere. Tests were made through an angle-of-attack range from -6 deg to 18 deg at dynamic pressures of 102 and 255 pounds per square foot with corresponding Mach numbers of 1.89 and 1.98 for the model both with and without a bumper arrangement designed to protect the rocket casing from the outer shell of the vehicle.
The Aerodynamic Drag of Flying-boat Hull Model as Measured in the NACA 20-foot Wind Tunnel I.
NASA Technical Reports Server (NTRS)
Hartman, Edwin P
1935-01-01
Measurements of aerodynamic drag were made in the 20-foot wind tunnel on a representative group of 11 flying-boat hull models. Four of the models were modified to investigate the effect of variations in over-all height, contours of deck, depth of step, angle of afterbody keel, and the addition of spray strips and windshields. The results of these tests, which cover a pitch-angle range from -5 to 10 degrees, are presented in a form suitable for use in performance calculations and for design purposes.
Some aspects of the aerodynamics of separating strap-ons
NASA Astrophysics Data System (ADS)
Biswas, K. K.; Krishnan, C. G.
1994-11-01
An aerodynamics model for analyzing strap-on separation is proposed. This model comprises both interference aerodynamics and free-body aerodynamics. The interference aerodynamics is primarily due to the close proximity of core and strap-ons. The free-body aerodynamics is solely due to the body geometry of the strap-ons. Using this aerodynamic model, the dynamics of separating strap-ons has been simulated in a six-degree-of-freedom mode to determine if a collision occurs. This aerodynamic model is very handy for various off-design studies relating to separating strap-ons.
Linearized aerodynamic and control law models of the X-29A airplane and comparison with flight data
NASA Technical Reports Server (NTRS)
Bosworth, John T.
1992-01-01
Flight control system design and analysis for aircraft rely on mathematical models of the vehicle dynamics. In addition to a six degree of freedom nonlinear simulation, the X-29A flight controls group developed a set of programs that calculate linear perturbation models throughout the X-29A flight envelope. The models include the aerodynamics as well as flight control system dynamics and were used for stability, controllability, and handling qualities analysis. These linear models were compared to flight test results to help provide a safe flight envelope expansion. A description is given of the linear models at three flight conditions and two flight control system modes. The models are presented with a level of detail that would allow the reader to reproduce the linear results if desired. Comparison between the response of the linear model and flight measured responses are presented to demonstrate the strengths and weaknesses of the linear models' ability to predict flight dynamics.
NASA Astrophysics Data System (ADS)
Ferreira, C.; Gonzalez, A.; Baldacchino, D.; Aparicio, M.; Gómez, S.; Munduate, X.; Garcia, N. R.; Sørensen, J. N.; Jost, E.; Knecht, S.; Lutz, T.; Chassapogiannis, P.; Diakakis, K.; Papadakis, G.; Voutsinas, S.; Prospathopoulos, J.; Gillebaart, T.; van Zuijlen, A.
2016-09-01
The FP7 AdVanced Aerodynamic Tools for lArge Rotors - Avatar project aims to develop and validate advanced aerodynamic models, to be used in integral design codes for the next generation of large scale wind turbines (10-20MW). One of the approaches towards reaching rotors for 10-20MW size is the application of flow control devices, such as flaps. In Task 3.2: Development of aerodynamic codes for modelling of flow devices on aerofoils and, rotors of the Avatar project, aerodynamic codes are benchmarked and validated against the experimental data of a DU95W180 airfoil in steady and unsteady flow, for different angle of attack and flap settings, including unsteady oscillatory trailing-edge-flap motion, carried out within the framework of WP3: Models for Flow Devices and Flow Control, Task 3.1: CFD and Experimental Database. The aerodynamics codes are: AdaptFoil2D, Foil2W, FLOWer, MaPFlow, OpenFOAM, Q3UIC, ATEFlap. The codes include unsteady Eulerian CFD simulations with grid deformation, panel models and indicial engineering models. The validation cases correspond to 18 steady flow cases, and 42 unsteady flow cases, for varying angle of attack, flap deflection and reduced frequency, with free and forced transition. The validation of the models show varying degrees of agreement, varying between models and flow cases.
Unsteady aerodynamics of blade rows
NASA Technical Reports Server (NTRS)
Verdon, Joseph M.
1989-01-01
The requirements placed on an unsteady aerodynamic theory intended for turbomachinery aeroelastic or aeroacoustic applications are discussed along with a brief description of the various theoretical models that are available to address these requirements. The major emphasis is placed on the description of a linearized inviscid theory which fully accounts for the affects of a nonuniform mean or steady flow on unsteady aerodynamic response. Although this linearization was developed primarily for blade flutter prediction, more general equations are presented which account for unsteady excitations due to incident external aerodynamic disturbances as well as those due to prescribed blade motions. The motivation for this linearized unsteady aerodynamic theory is focused on, its physical and mathematical formulation is outlined and examples are presented to illustrate the status of numerical solution procedures and several effects of mean flow nonuniformity on unsteady aerodynamic response.
Active Aeroelastic Wing Aerodynamic Model Development and Validation for a Modified F/A-18A Airplane
NASA Technical Reports Server (NTRS)
Cumming, Stephen B.; Diebler, Corey G.
2005-01-01
A new aerodynamic model has been developed and validated for a modified F/A-18A airplane used for the Active Aeroelastic Wing (AAW) research program. The goal of the program was to demonstrate the advantages of using the inherent flexibility of an aircraft to enhance its performance. The research airplane was an F/A-18A with wings modified to reduce stiffness and a new control system to increase control authority. There have been two flight phases. Data gathered from the first flight phase were used to create the new aerodynamic model. A maximum-likelihood output-error parameter estimation technique was used to obtain stability and control derivatives. The derivatives were incorporated into the National Aeronautics and Space Administration F-18 simulation, validated, and used to develop new AAW control laws. The second phase of flights was used to evaluate the handling qualities of the AAW airplane and the control law design process, and to further test the accuracy of the new model. The flight test envelope covered Mach numbers between 0.85 and 1.30 and dynamic pressures from 600 to 1250 pound-force per square foot. The results presented in this report demonstrate that a thorough parameter identification analysis can be used to improve upon models that were developed using other means. This report describes the parameter estimation technique used, details the validation techniques, discusses differences between previously existing F/A-18 models, and presents results from the second phase of research flights.
A program to evaluate a control system based on feedback of aerodynamic pressure differentials
NASA Technical Reports Server (NTRS)
Levy, D. W.; Finn, P.; Roskam, J.
1981-01-01
The use of aerodynamic pressure differentials to position a control surface is evaluated. The system is a differential pressure command loop, analogous to a position command loop, where the surface is commanded to move until a desired differential pressure across the surface is achieved. This type of control is more direct and accurate because it is the differential pressure which causes the control forces and moments. A frequency response test was performed in a low speed wind tunnel to measure the performance of the system. Both pressure and position feedback were tested. The pressure feedback performed as well as position feedback implying that the actuator, with a break frequency on the order of 10 Rad/sec, was the limiting component. Theoretical considerations indicate that aerodynamic lags will not appear below frequencies of 50 Rad/sec, or higher.
NASA Astrophysics Data System (ADS)
Amiraux, Mathieu
Rotorcraft Blade-Vortex Interaction (BVI) remains one of the most challenging flow phenomenon to simulate numerically. Over the past decade, the HART-II rotor test and its extensive experimental dataset has been a major database for validation of CFD codes. Its strong BVI signature, with high levels of intrusive noise and vibrations, makes it a difficult test for computational methods. The main challenge is to accurately capture and preserve the vortices which interact with the rotor, while predicting correct blade deformations and loading. This doctoral dissertation presents the application of a coupled CFD/CSD methodology to the problem of helicopter BVI and compares three levels of fidelity for aerodynamic modeling: a hybrid lifting-line/free-wake (wake coupling) method, with modified compressible unsteady model; a hybrid URANS/free-wake method; and a URANS-based wake capturing method, using multiple overset meshes to capture the entire flow field. To further increase numerical correlation, three helicopter fuselage models are implemented in the framework. The first is a high resolution 3D GPU panel code; the second is an immersed boundary based method, with 3D elliptic grid adaption; the last one uses a body-fitted, curvilinear fuselage mesh. The main contribution of this work is the implementation and systematic comparison of multiple numerical methods to perform BVI modeling. The trade-offs between solution accuracy and computational cost are highlighted for the different approaches. Various improvements have been made to each code to enhance physical fidelity, while advanced technologies, such as GPU computing, have been employed to increase efficiency. The resulting numerical setup covers all aspects of the simulation creating a truly multi-fidelity and multi-physics framework. Overall, the wake capturing approach showed the best BVI phasing correlation and good blade deflection predictions, with slightly under-predicted aerodynamic loading magnitudes
Improved Aerodynamic Analysis for Hybrid Wing Body Conceptual Design Optimization
NASA Technical Reports Server (NTRS)
Gern, Frank H.
2012-01-01
This paper provides an overview of ongoing efforts to develop, evaluate, and validate different tools for improved aerodynamic modeling and systems analysis of Hybrid Wing Body (HWB) aircraft configurations. Results are being presented for the evaluation of different aerodynamic tools including panel methods, enhanced panel methods with viscous drag prediction, and computational fluid dynamics. Emphasis is placed on proper prediction of aerodynamic loads for structural sizing as well as viscous drag prediction to develop drag polars for HWB conceptual design optimization. Data from transonic wind tunnel tests at the Arnold Engineering Development Center s 16-Foot Transonic Tunnel was used as a reference data set in order to evaluate the accuracy of the aerodynamic tools. Triangularized surface data and Vehicle Sketch Pad (VSP) models of an X-48B 2% scale wind tunnel model were used to generate input and model files for the different analysis tools. In support of ongoing HWB scaling studies within the NASA Environmentally Responsible Aviation (ERA) program, an improved finite element based structural analysis and weight estimation tool for HWB center bodies is currently under development. Aerodynamic results from these analyses are used to provide additional aerodynamic validation data.
1975-11-01
further improve the contrast all of the interior surfaces of the test chamber are painted flat black and the bac!-,ground walls in view of the cameras...to be adequate to eliminate wall effects on the chaff aerodynamics. Secondly, the chamber air mass had to be sufficiently small that it would damp out...independently- supported special rotating-shutter system to "strobe" the dipole images. The integral shutter in each lens assembly is also retained for
1980-06-01
The perturbation velozity components are due to the blade degrees of freedom, the shaft motion, and the aerodynamic gust velocity: -tA 4- + V(. o ) -A...gimballed, and teetering rotors with an arbitrary number of blades. The rotor degrees of freedom included are blade flap/lag bending, rigid pitch and elastic...tunnel is also covered. The aircraft degrees of freedom included are the six rigid body motions, elastic airframe motions, and the rotor/engine speed
NASA Technical Reports Server (NTRS)
Kuhlman, J. M.
1983-01-01
Wind tunnel test results have been presented herein for a subsonic transport type wing fitted with winglets. Wind planform was chosen to be representative of wings used on current jet transport aircraft, while wing and winglet camber surfaces were designed using two different linear aerodynamic design methods. The purpose of the wind tunnel investigation was to determine the effectiveness of these linear aerodynamic design computer codes in designing a non-planar transport configuration which would cruise efficiently. The design lift coefficient was chosen to be 0.4, at a design Mach number of 0.8. Force and limited pressure data were obtained for the basic wing, and for the wing fitted with the two different winglet designs, at Mach numbers of 0.60, 0.70, 0.75 and 0.80 over an angle of attack range of -2 to +6 degrees, at zero sideslip. The data have been presented without analysis to expedite publication.
NASA Technical Reports Server (NTRS)
Pototzky, Anthony S.
2008-01-01
A simple matrix polynomial approach is introduced for approximating unsteady aerodynamics in the s-plane and ultimately, after combining matrix polynomial coefficients with matrices defining the structure, a matrix polynomial of the flutter equations of motion (EOM) is formed. A technique of recasting the matrix-polynomial form of the flutter EOM into a first order form is also presented that can be used to determine the eigenvalues near the origin and everywhere on the complex plane. An aeroservoelastic (ASE) EOM have been generalized to include the gust terms on the right-hand side. The reasons for developing the new matrix polynomial approach are also presented, which are the following: first, the "workhorse" methods such as the NASTRAN flutter analysis lack the capability to consistently find roots near the origin, along the real axis or accurately find roots farther away from the imaginary axis of the complex plane; and, second, the existing s-plane methods, such as the Roger s s-plane approximation method as implemented in ISAC, do not always give suitable fits of some tabular data of the unsteady aerodynamics. A method available in MATLAB is introduced that will accurately fit generalized aerodynamic force (GAF) coefficients in a tabular data form into the coefficients of a matrix polynomial form. The root-locus results from the NASTRAN pknl flutter analysis, the ISAC-Roger's s-plane method and the present matrix polynomial method are presented and compared for accuracy and for the number and locations of roots.
NASA Technical Reports Server (NTRS)
Nelson, D. P.; Morris, P. M.
1980-01-01
Aerodynamic performance and jet noise characteristics of a one sixth scale model of the variable cycle engine testbed exhaust system were obtained in a series of static tests over a range of simulated engine operating conditions. Model acoustic data were acquired. Data were compared to predictions of coannular model nozzle performance. The model, tested with an without a hardwall ejector, had a total flow area equivalent to a 0.127 meter (5 inch) diameter conical nozzle with a 0.65 fan to primary nozzle area ratio and a 0.82 fan nozzle radius ratio. Fan stream temperatures and velocities were varied from 422 K to 1089 K (760 R to 1960 R) and 434 to 755 meters per second (1423 to 2477 feet per second). Primary stream properties were varied from 589 to 1089 K (1060 R to 1960 R) and 353 to 600 meters per second (1158 to 1968 feet per second). Exhaust plume velocity surveys were conducted at one operating condition with and without the ejector installed. Thirty aerodynamic performance data points were obtained with an unheated air supply. Fan nozzle pressure ratio was varied from 1.8 to 3.2 at a constant primary pressure ratio of 1.6; primary pressure ratio was varied from 1.4 to 2.4 while holding fan pressure ratio constant at 2.4. Operation with the ejector increased nozzle thrust coefficient 0.2 to 0.4 percent.
Payload vehicle aerodynamic reentry analysis
NASA Astrophysics Data System (ADS)
Tong, Donald
An approach for analyzing the dynamic behavior of a cone-cylinder payload vehicle during reentry to insure proper deployment of the parachute system and recovery of the payload is presented. This analysis includes the study of an aerodynamic device that is useful in extending vehicle axial rotation through the maximum dynamic pressure region. Attention is given to vehicle configuration and reentry trajectory, the derivation of pitch static aerodynamics, the derivation of the pitch damping coefficient, pitching moment modeling, aerodynamic roll device modeling, and payload vehicle reentry dynamics. It is shown that the vehicle dynamics at parachute deployment are well within the design limit of the recovery system, thus ensuring successful payload recovery.
Aerodynamic preliminary analysis system 2. Part 1: Theory
NASA Technical Reports Server (NTRS)
Bonner, E.; Clever, W.; Dunn, K.
1981-01-01
A subsonic/supersonic/hypersonic aerodynamic analysis was developed by integrating the Aerodynamic Preliminary Analysis System (APAS), and the inviscid force calculation modules of the Hypersonic Arbitrary Body Program. APAS analysis was extended for nonlinear vortex forces using a generalization of the Polhamus analogy. The interactive system provides appropriate aerodynamic models for a single input geometry data base and has a run/output format similar to a wind tunnel test program. The user's manual was organized to cover the principle system activities of a typical application, geometric input/editing, aerodynamic evaluation, and post analysis review/display. Sample sessions are included to illustrate the specific task involved and are followed by a comprehensive command/subcommand dictionary used to operate the system.
Computational Aerodynamic Analysis of Offshore Upwind and Downwind Turbines
Zhao, Qiuying; Sheng, Chunhua; Afjeh, Abdollah
2014-01-01
Aerodynamic interactions of the model NREL 5 MW offshore horizontal axis wind turbines (HAWT) are investigated using a high-fidelity computational fluid dynamics (CFD) analysis. Four wind turbine configurations are considered; three-bladed upwind and downwind and two-bladed upwind and downwind configurations, which operate at two different rotor speeds of 12.1 and 16 RPM. In the present study, both steady and unsteady aerodynamic loads, such as the rotor torque, blade hub bending moment, and base the tower bending moment of the tower, are evaluated in detail to provide overall assessment of different wind turbine configurations. Aerodynamic interactions between the rotor and tower are analyzed,more » including the rotor wake development downstream. The computational analysis provides insight into aerodynamic performance of the upwind and downwind, two- and three-bladed horizontal axis wind turbines.« less
Improved two-equation k-omega turbulence models for aerodynamic flows
NASA Technical Reports Server (NTRS)
Menter, Florian R.
1992-01-01
Two new versions of the k-omega two-equation turbulence model will be presented. The new Baseline (BSL) model is designed to give results similar to those of the original k-omega model of Wilcox, but without its strong dependency on arbitrary freestream values. The BSL model is identical to the Wilcox model in the inner 50 percent of the boundary-layer but changes gradually to the high Reynolds number Jones-Launder k-epsilon model (in a k-omega formulation) towards the boundary-layer edge. The new model is also virtually identical to the Jones-Lauder model for free shear layers. The second version of the model is called Shear-Stress Transport (SST) model. It is based on the BSL model, but has the additional ability to account for the transport of the principal shear stress in adverse pressure gradient boundary-layers. The model is based on Bradshaw's assumption that the principal shear stress is proportional to the turbulent kinetic energy, which is introduced into the definition of the eddy-viscosity. Both models are tested for a large number of different flowfields. The results of the BSL model are similar to those of the original k-omega model, but without the undesirable freestream dependency. The predictions of the SST model are also independent of the freestream values and show excellent agreement with experimental data for adverse pressure gradient boundary-layer flows.
NASA Astrophysics Data System (ADS)
Krasinsky, D. V.; Salomatov, V. V.; Anufriev, I. S.; Sharypov, O. V.; Shadrin, E. Yu.; Anikin, Yu. A.
2015-02-01
Some results of the complex experimental and numerical study of aerodynamics and transfer processes in a vortex furnace, whose design was improved via the distributed tangential injection of fuel-air flows through the upper and lower burners, were presented. The experimental study of the aerodynamic characteristics of a spatial turbulent flow was performed on the isothermal laboratory model (at a scale of 1 : 20) of an improved vortex furnace using a laser Doppler measurement system. The comparison of experimental data with the results of the numerical modeling of an isothermal flow for the same laboratory furnace model demonstrated their agreement to be acceptable for engineering practice.
NASA Technical Reports Server (NTRS)
Abdol-Hamid, Khaled S.; Ghaffari, Farhad
2011-01-01
Numerical predictions of the longitudinal aerodynamic characteristics for the Ares I class of vehicles, along with the associated error estimate derived from an iterative convergence grid refinement, are presented. Computational results are based on the unstructured grid, Reynolds-averaged Navier-Stokes flow solver USM3D, with an assumption that the flow is fully turbulent over the entire vehicle. This effort was designed to complement the prior computational activities conducted over the past five years in support of the Ares I Project with the emphasis on the vehicle s last design cycle designated as the A106 configuration. Due to a lack of flight data for this particular design s outer mold line, the initial vehicle s aerodynamic predictions and the associated error estimates were first assessed and validated against the available experimental data at representative wind tunnel flow conditions pertinent to the ascent phase of the trajectory without including any propulsion effects. Subsequently, the established procedures were then applied to obtain the longitudinal aerodynamic predictions at the selected flight flow conditions. Sample computed results and the correlations with the experimental measurements are presented. In addition, the present analysis includes the relevant data to highlight the balance between the prediction accuracy against the grid size and, thus, the corresponding computer resource requirements for the computations at both wind tunnel and flight flow conditions. NOTE: Some details have been removed from selected plots and figures in compliance with the sensitive but unclassified (SBU) restrictions. However, the content still conveys the merits of the technical approach and the relevant results.
NASA Technical Reports Server (NTRS)
Jacobs, P. F.
1985-01-01
An investigation was conducted in the Langley 8 Foot Transonic Pressure Tunnel to determine the effect of aileron deflections on the aerodynamic characteristics of a subsonic energy efficient transport (EET) model. The semispan model had an aspect ratio 10 supercritical wing and was configured with a conventionally located set of ailerons (i.e., a high speed aileron located inboard and a low speed aileron located outboard). Data for the model were taken over a Mach number range from 0.30 to 0.90 and an angle of attack range from approximately -2 deg to 10 deg. The Reynolds number was 2.5 million per foot for Mach number = 0.30 and 4 million per foot for the other Mach numbers. Model force and moment data, aileron effectiveness parameters, aileron hinge moment data, otherwise pressure distributions, and spanwise load data are presented.
NASA Technical Reports Server (NTRS)
Persing, T. Ray; Bellish, Christine A.; Brandon, Jay; Kenney, P. Sean; Carzoo, Susan; Buttrill, Catherine; Guenther, Arlene
2005-01-01
Several aircraft airframe modeling approaches are currently being used in the DoD community for acquisition, threat evaluation, training, and other purposes. To date there has been no clear empirical study of the impact of airframe simulation fidelity on piloted real-time aircraft simulation study results, or when use of a particular level of fidelity is indicated. This paper documents a series of piloted simulation studies using three different levels of airframe model fidelity. This study was conducted using the NASA Langley Differential Maneuvering Simulator. Evaluations were conducted with three pilots for scenarios requiring extensive maneuvering of the airplanes during air combat. In many cases, a low-fidelity modified point-mass model may be sufficient to evaluate the combat effectiveness of the aircraft. However, in cases where high angle-of-attack flying qualities and aerodynamic performance are a factor or when precision tracking ability of the aircraft must be represented, use of high-fidelity models is indicated.
On Cup Anemometer Rotor Aerodynamics
Pindado, Santiago; Pérez, Javier; Avila-Sanchez, Sergio
2012-01-01
The influence of anemometer rotor shape parameters, such as the cups' front area or their center rotation radius on the anemometer's performance was analyzed. This analysis was based on calibrations performed on two different anemometers (one based on magnet system output signal, and the other one based on an opto-electronic system output signal), tested with 21 different rotors. The results were compared to the ones resulting from classical analytical models. The results clearly showed a linear dependency of both calibration constants, the slope and the offset, on the cups' center rotation radius, the influence of the front area of the cups also being observed. The analytical model of Kondo et al. was proved to be accurate if it is based on precise data related to the aerodynamic behavior of a rotor's cup. PMID:22778638
On cup anemometer rotor aerodynamics.
Pindado, Santiago; Pérez, Javier; Avila-Sanchez, Sergio
2012-01-01
The influence of anemometer rotor shape parameters, such as the cups' front area or their center rotation radius on the anemometer's performance was analyzed. This analysis was based on calibrations performed on two different anemometers (one based on magnet system output signal, and the other one based on an opto-electronic system output signal), tested with 21 different rotors. The results were compared to the ones resulting from classical analytical models. The results clearly showed a linear dependency of both calibration constants, the slope and the offset, on the cups' center rotation radius, the influence of the front area of the cups also being observed. The analytical model of Kondo et al. was proved to be accurate if it is based on precise data related to the aerodynamic behavior of a rotor's cup.
Zonal Two Equation Kappa-Omega Turbulence Models for Aerodynamic Flows
NASA Technical Reports Server (NTRS)
Menter, Florian R.
1993-01-01
Two new versions of the kappa-omega two-equation turbulence model will be presented. The new Baseline (BSL) model is designed to give results similar to those of the original kappa-omega model of Wilcox, but without its strong dependency on arbitrary freestream values. The BSL model is identical to the Wilcox model in the inner 50% of the boundary-layer but changes gradually to the standard kappa-epsilon model (in a kappa- omega formulation) towards the boundary-layer edge. The free shear layers. The second version of the model is called Shear-Stress Transport (SST) model. It is a variation of the BSL model with the additional ability to account for the transport of the principal turbulent shear stress in adverse pressure gradient boundary-layers. The model is based on Bradshaw's assumption that the principal shear-stress is proportional to the turbulent kinetic energy, which is introduced into the definition of the eddy-viscosity. Both models are tested for a large number of different flowfields. The results of the BSL model are similar to those of the original kappa-omega model, but without the undesirable freestream dependency. The predictions of the SST model are also independent of the freestream values but show better agreement with experimental data for adverse pressure gradient boundary-layer flows.
NASA Astrophysics Data System (ADS)
Toyota, Kenjiro; Dastoor, Ashu P.; Ryzhkov, Andrei
2016-12-01
Turbulence controls the vertical transfer of momentum, heat and trace constituents in the atmospheric boundary layer. In the lowest 10% of this layer lies the surface boundary layer (SBL) where the vertical fluxes of transferred quantities have been successfully parameterized using the Monin-Obukhov similarity theory in weather forecast, climate and atmospheric chemistry models. However, there is a large degree of empiricism in the stability-correction parameterizations used to formulate eddy diffusivity and aerodynamic resistance particularly under strongly stable ambient conditions. Although the influence of uncertainties in stability-correction parameterizations on eddy diffusivity is actively studied in boundary-layer meteorological modeling, its impact on dry deposition in atmospheric chemistry modeling is not well characterized. In this study, we address this gap by providing the mathematical basis for the relationship between the formulations of vertical surface flux used in meteorological and atmospheric chemistry modeling communities, and by examining the sensitivity of the modeled dry deposition velocities in statically stable SBL to the choice of stability-correction parameterizations used in three operational and/or research environmental models (GEM/GEM-MACH, ECMWF IFS and CMAQ-MM5). Aerodynamic resistances (ra) calculated by the three sets of parameterizations are notably different from each other and are also different from those calculated by a "z-less" scaling formulation under strongly stable conditions (the bulk Richardson number > 0.2). Furthermore, we show that many atmospheric chemistry models calculate ra using formulations which are inconsistent with the derivation of micro-meteorological parameters. Finally, practical implications of the differences in stability-correction algorithms are discussed for the computations of dry deposition velocities of SO2, O3 and reactive bromine compounds for specific cases of stable SBL.
NASA Technical Reports Server (NTRS)
Dawson, John R; Hartman, Edwin P
1938-01-01
Four models of outboard floats (N.A.C.A. models 51-A, 51-B, 51-C, and 51-D) were tested in the N.A.C.A. tank to determine their hydrodynamic characteristics and in the 20-foot wind tunnel to determine their aerodynamic drag. The results of the tests, together with comparisons of them, are presented in the form of charts. From the comparisons, the order of merit of the models is estimated for each factor considered. The best compromise between the various factors seems to be given by model 51-D. This model is the only one in the series with a transverse step.
NASA Technical Reports Server (NTRS)
Nelson, D. P.
1983-01-01
Wind tunnel model tests were conducted to demonstrate the aerodynamic performance improvements of a refined actuated inlet ejector nozzle. Models of approximately one-tenth scale were configured to simulate nozzle operation at takeoff, subsonic cruise, transonic cruise and supersonic cruise. Variations of model components provided a performance evaluation of ejector inlet and exit area, forebody boattail angle and ejector inlet operation in the open and closed mode. Approximately 700 data points were acquired at Mach numbers of 0, 0.36, 0.9, 1.2, and 2.0 for a wide range of nozzle flow conditions. Results show that relative to two ejector nozzles previously tested performance was improved significantly at takeoff and subsonic cruise performance, a C sub f of 0.982, was attained equal to the high performance of the previous tests. The established advanced supersonic transport propulsion study performance goals were met or closely approached at takeoff and supersonic cruise.
Aerodynamic Reconstruction Applied to Parachute Test Vehicle Flight Data Analysis
NASA Technical Reports Server (NTRS)
Cassady, Leonard D.; Ray, Eric S.; Truong, Tuan H.
2013-01-01
The aerodynamics, both static and dynamic, of a test vehicle are critical to determining the performance of the parachute cluster in a drop test and for conducting a successful test. The Capsule Parachute Assembly System (CPAS) project is conducting tests of NASA's Orion Multi-Purpose Crew Vehicle (MPCV) parachutes at the Army Yuma Proving Ground utilizing the Parachute Test Vehicle (PTV). The PTV shape is based on the MPCV, but the height has been reduced in order to fit within the C-17 aircraft for extraction. Therefore, the aerodynamics of the PTV are similar, but not the same as, the MPCV. A small series of wind tunnel tests and computational fluid dynamics cases were run to modify the MPCV aerodynamic database for the PTV, but aerodynamic reconstruction of the flights has proven an effective source for further improvements to the database. The acceleration and rotational rates measured during free flight, before parachute inflation but during deployment, were used to con rm vehicle static aerodynamics. A multibody simulation is utilized to reconstruct the parachute portions of the flight. Aerodynamic or parachute parameters are adjusted in the simulation until the prediction reasonably matches the flight trajectory. Knowledge of the static aerodynamics is critical in the CPAS project because the parachute riser load measurements are scaled based on forebody drag. PTV dynamic damping is critical because the vehicle has no reaction control system to maintain attitude - the vehicle dynamics must be understood and modeled correctly before flight. It will be shown here that aerodynamic reconstruction has successfully contributed to the CPAS project.
Experimental and theoretical aerodynamic characteristics of a high-lift semispan wing model
NASA Technical Reports Server (NTRS)
Applin, Zachary T.; Gentry, Garl L., Jr.
1990-01-01
Experimental and theoretical aerodynamic characteristics were compared for a high-lift, semispan wing configuration that incorporated a slightly modified version of the NASA Advanced Laminar Flow Control airfoil section. The experimental investigation was conducted in the Langley 14- by 22-Foot Subsonic Tunnel at chord Reynolds numbers of 2.36 and 3.33 million. A two-dimensional airfoil code and a three-dimensional panel code were used to obtain aerodynamic predictions. Two-dimensional data were corrected for three-dimensional effects. Comparisons between predicted and measured values were made for the cruise configuration and for various high-lift configurations. Both codes predicted lift and pitching moment coefficients that agreed well with experiment for the cruise configuration. These parameters were overpredicted for all high-lift configurations. Drag coefficient was underpredicted for all cases. Corrected two-dimensional pressure distributions typically agreed well with experiment, while the panel code overpredicted the leading-edge suction peak on the wing. One important feature missing from both of these codes was a capability for separated flow analysis. The major cause of disparity between the measured data and predictions presented herein was attributed to separated flow conditions.
Modeling and Simulation of Radiative Compressible Flows in Aerodynamic Heating Arc-Jet Facility
NASA Technical Reports Server (NTRS)
Bensassi, Khalil; Laguna, Alejandro A.; Lani, Andrea; Mansour, Nagi N.
2016-01-01
Numerical simulations of an arc heated flow inside NASA's 20 [MW] Aerodynamics heating facility (AHF) are performed in order to investigate the three-dimensional swirling flow and the current distribution inside the wind tunnel. The plasma is considered in Local Thermodynamics Equilibrium(LTE) and is composed of Air-Argon gas mixture. The governing equations are the Navier-Stokes equations that include source terms corresponding to Joule heating and radiative cooling. The former is obtained by solving an electric potential equation, while the latter is calculated using an innovative massively parallel ray-tracing algorithm. The fully coupled system is closed by the thermodynamics relations and transport properties which are obtained from Chapman-Enskog method. A novel strategy was developed in order to enable the flow solver and the radiation calculation to be preformed independently and simultaneously using a different number of processors. Drastic reduction in the computational cost was achieved using this strategy. Details on the numerical methods used for space discretization, time integration and ray-tracing algorithm will be presented. The effect of the radiative cooling on the dynamics of the flow will be investigated. The complete set of equations were implemented within the COOLFluiD Framework. Fig. 1 shows the geometry of the Anode and part of the constrictor of the Aerodynamics heating facility (AHF). Fig. 2 shows the velocity field distribution along (x-y) plane and the streamline in (z-y) plane.
NASA Technical Reports Server (NTRS)
Sulyma, P. R.; Penny, M. M.
1978-01-01
A base pressure data correlation study was conducted to define exhaust plume similarity parameters for use in Space Shuttle power-on launch vehicle aerodynamic test programs. Data correlations were performed for single bodies having, respectively, single and triple nozzle configurations and for a triple body configuration with single nozzles on each of the outside bodies. Base pressure similarity parameters were found to differ for the single nozzle and triple nozzle configurations. However, the correlation parameter for each was found to be a strong function of the nozzle exit momentum. Results of the data base evaluation are presented indicating an assessment of all data points. Analytical/experimental data comparisons were made for nozzle calibrations and correction factors derived, where indicated for use in nozzle exit plane data calculations.
The aerodynamic and acoustic characteristics of an over-the-wing target-type thrust reverser model
NASA Technical Reports Server (NTRS)
Falarski, M. D.
1976-01-01
A static test of a large-scale, over-the-wing (OTW) powered-lift model was performed. The OTW propulsion system had been modified to incorporate a simple target-type thrust reverser as well as the normal rectangular OTW exhaust nozzle. Tests were performed in both the reverse thrust and approach configurations. The thrust reverser noise created by jet turbulence mixing and the OTW approach noise were both low frequency and broadband. When scaled to a 45,400-kg (100,000-lb) aircraft, the thrust reverser and approach configurations produced peak 152-m (500-ft) sideline perceived noise levels of 110 and 105 PNdB, respectively. The aerodynamic performance of the model showed that 50% or greater reverser effectiveness can be achieved without experiencing ingestion of exhaust gas or ground debris into the engine inlets.
NASA Technical Reports Server (NTRS)
Brandon, Jay M.; Shah, Gautam H.
1990-01-01
The effects of harmonic or constant-rate-ramp pitching motions (giving angles of attack from 0 to 75 deg) on the aerodynamic performance of a fighter-aircraft model with highly swept leading-edge extensions are investigated experimentally in the NASA Langley 12-ft low-speed wind tunnel. The model configuration and experimental setup are described, and the results of force and moment measurements and flow visualizations are presented graphically and discussed in detail. Large force overshoots and hysteresis are observed and attributed to lags in vortical-flow development and breakup. The motion variables have a strong influence on the persistence of dynamic effects, which are found to affect pitch-rate capability more than flight-path turning performance.
NASA Technical Reports Server (NTRS)
Frassinelli, Mark C.; Carson, George T., Jr.
1990-01-01
An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of horizontal and vertical tail size reductions on the longitudinal aerodynamic characteristics of a modified F-15 model with canards and 2-D convergent-divergent nozzles. Quantifying the drag decrease at low angles of attack produced by tail size reductions was the primary focus. The model was tested at Mach numbers of 0.40, 0.90, and 1.20 over an angle of attack of -2 degree to 10 degree. The nozzle exhaust flow was simulated using high pressure air at nozzle pressure ratios varying from 1.0 (jet off) to 7.5. Data were obtained on the baseline configuration with and without tails as well as with reduced horizontal and/or vertical tail sizes that were 75, 50, and 25 percent of the baseline tail areas.
NASA Technical Reports Server (NTRS)
Stromberg, W. J.
1981-01-01
An engine was specially prepared with extensive instrumentation to monitor performance, case temperatures, and clearance changes. A special loading device was used to apply known loads on the engine by the use of cables placed around the flight inlet. These loads simulated the estimated aerodynamic pressure distributions that occur on the inlet in various segments of a typical airplane flight. Test results indicate that the engine lost 1.3 percent in take-off thrust specific fuel consumption (TSFC) during the course of the test effort. Permanent clearance changes due to the loads accounted for 1.1 percent; increase in low pressure compressor airfoil roughness and thermal distortion in the high pressure turbine accounted for 0.2 percent. Pretest predicted performance loss due to clearance changes was 0.9 percent in TSFC. Therefore, the agreement between measurement and prediction is considered to be excellent.
Wang, Ji Kang; Sun, Mao
2005-10-01
The aerodynamics and forewing-hindwing interaction of a model dragonfly in forward flight are studied, using the method of numerically solving the Navier-Stokes equations. Available morphological and stroke-kinematic parameters of dragonfly (Aeshna juncea) are used for the model dragonfly. Six advance ratios (J; ranging from 0 to 0.75) and, at each J, four forewing-hindwing phase angle differences (gamma(d); 180 degrees, 90 degrees, 60 degrees and 0 degree) are considered. The mean vertical force and thrust are made to balance the weight and body-drag, respectively, by adjusting the angles of attack of the wings, so that the flight could better approximate the real flight. At hovering and low J (J=0, 0.15), the model dragonfly uses separated flows or leading-edge vortices (LEV) on both the fore- and hindwing downstrokes; at medium J (J=0.30, 0.45), it uses the LEV on the forewing downstroke and attached flow on the hindwing downstroke; at high J (J=0.6, 0.75), it uses attached flows on both fore- and hindwing downstrokes. (The upstrokes are very lightly loaded and, in general, the flows are attached.) At a given J, at gamma(d)=180 degrees, there are two vertical force peaks in a cycle, one in the first half of the cycle, produced mainly by the hindwing downstroke, and the other in the second half of the cycle, produced mainly by the forewing downstroke; at gamma(d)=90 degrees, 60 degrees and 0 degree, the two force peaks merge into one peak. The vertical force is close to the resultant aerodynamic force [because the thrust (or body-drag) is much smaller than vertical force (or the weight)]. 55-65% of the vertical force is contributed by the drag of the wings. The forewing-hindwing interaction is detrimental to the vertical force (and resultant force) generation. At hovering, the interaction reduces the mean vertical force (and resultant force) by 8-15%, compared with that without interaction; as J increases, the reduction generally decreases (e.g. at J=0.6 and
A CFD-based aerodynamic design procedure for hypersonic wind-tunnel nozzles
NASA Technical Reports Server (NTRS)
Korte, John J.
1993-01-01
A new procedure which unifies the best of current classical design practices, computational fluid dynamics (CFD), and optimization procedures is demonstrated for designing the aerodynamic lines of hypersonic wind-tunnel nozzles. The new procedure can be used to design hypersonic wind tunnel nozzles with thick boundary layers where the classical design procedure has been shown to break down. An efficient CFD code, which solves the parabolized Navier-Stokes (PNS) equations using an explicit upwind algorithm, is coupled to a least-squares (LS) optimization procedure. A LS problem is formulated to minimize the difference between the computed flow field and the objective function, consisting of the centerline Mach number distribution and the exit Mach number and flow angle profiles. The aerodynamic lines of the nozzle are defined using a cubic spline, the slopes of which are optimized with the design procedure. The advantages of the new procedure are that it allows full use of powerful CFD codes in the design process, solves an optimization problem to determine the new contour, can be used to design new nozzles or improve sections of existing nozzles, and automatically compensates the nozzle contour for viscous effects as part of the unified design procedure. The new procedure is demonstrated by designing two Mach 15, a Mach 12, and a Mach 18 helium nozzles. The flexibility of the procedure is demonstrated by designing the two Mach 15 nozzles using different constraints, the first nozzle for a fixed length and exit diameter and the second nozzle for a fixed length and throat diameter. The computed flow field for the Mach 15 least squares parabolized Navier-Stokes (LS/PNS) designed nozzle is compared with the classically designed nozzle and demonstrates a significant improvement in the flow expansion process and uniform core region.
Advanced Response Surface Modeling of Ares I Roll Control Jet Aerodynamic Interactions
NASA Technical Reports Server (NTRS)
Favaregh, Noah M.
2010-01-01
The Ares I rocket uses roll control jets. These jets have aerodynamic implications as they impinge on the surface and protuberances of the vehicle. The jet interaction on the body can cause an amplification or a reduction of the rolling moment produced by the jet itself, either increasing the jet effectiveness or creating an adverse effect. A design of experiments test was planned and carried out using computation fluid dynamics, and a subsequent response surface analysis ensued on the available data to characterize the jet interaction across the ascent portion of the Ares I flight envelope. Four response surface schemes were compared including a single response surface covering the entire design space, separate sector responses that did not overlap, continuously overlapping surfaces, and recursive weighted response surfaces. These surfaces were evaluated on traditional statistical metrics as well as visual inspection. Validation of the recursive weighted response surface was performed using additionally available data at off-design point locations.
NASA Technical Reports Server (NTRS)
Hughes, Christopher E.; Podboy, Gary, G.; Woodward, Richard P.; Jeracki, Robert, J.
2013-01-01
The design of effective new technologies to reduce aircraft propulsion noise is dependent on identifying and understanding the noise sources and noise generation mechanisms in the modern turbofan engine, as well as determining their contribution to the overall aircraft noise signature. Therefore, a comprehensive aeroacoustic wind tunnel test program was conducted called the Fan Broadband Source Diagnostic Test as part of the NASA Quiet Aircraft Technology program. The test was performed in the anechoic NASA Glenn 9- by 15-Foot Low Speed Wind Tunnel using a 1/5 scale model turbofan simulator which represented a current generation, medium pressure ratio, high bypass turbofan aircraft engine. The investigation focused on simulating in model scale only the bypass section of the turbofan engine. The test objectives were to: identify the noise sources within the model and determine their noise level; investigate several component design technologies by determining their impact on the aerodynamic and acoustic performance of the fan stage; and conduct detailed flow diagnostics within the fan flow field to characterize the physics of the noise generation mechanisms in a turbofan model. This report discusses results obtained for one aspect of the Source Diagnostic Test that investigated the effect of the bypass or fan nozzle exit area on the bypass stage aerodynamic performance, specifically the fan and outlet guide vanes or stators, as well as the farfield acoustic noise level. The aerodynamic performance, farfield acoustics, and Laser Doppler Velocimeter flow diagnostic results are presented for the fan and four different fixed-area bypass nozzle configurations. The nozzles simulated fixed engine operating lines and encompassed the fan stage operating envelope from near stall to cruise. One nozzle was selected as a baseline reference, representing the nozzle area which would achieve the design point operating conditions and fan stage performance. The total area change from
Computational aerodynamics and artificial intelligence
NASA Technical Reports Server (NTRS)
Mehta, U. B.; Kutler, P.
1984-01-01
The general principles of artificial intelligence are reviewed and speculations are made concerning how knowledge based systems can accelerate the process of acquiring new knowledge in aerodynamics, how computational fluid dynamics may use expert systems, and how expert systems may speed the design and development process. In addition, the anatomy of an idealized expert system called AERODYNAMICIST is discussed. Resource requirements for using artificial intelligence in computational fluid dynamics and aerodynamics are examined. Three main conclusions are presented. First, there are two related aspects of computational aerodynamics: reasoning and calculating. Second, a substantial portion of reasoning can be achieved with artificial intelligence. It offers the opportunity of using computers as reasoning machines to set the stage for efficient calculating. Third, expert systems are likely to be new assets of institutions involved in aeronautics for various tasks of computational aerodynamics.
The aerodynamics of hovering flight in Drosophila.
Fry, Steven N; Sayaman, Rosalyn; Dickinson, Michael H
2005-06-01
Using 3D infrared high-speed video, we captured the continuous wing and body kinematics of free-flying fruit flies, Drosophila melanogaster, during hovering and slow forward flight. We then 'replayed' the wing kinematics on a dynamically scaled robotic model to measure the aerodynamic forces produced by the wings. Hovering animals generate a U-shaped wing trajectory, in which large drag forces during a downward plunge at the start of each stroke create peak vertical forces. Quasi-steady mechanisms could account for nearly all of the mean measured force required to hover, although temporal discrepancies between instantaneous measured forces and model predictions indicate that unsteady mechanisms also play a significant role. We analyzed the requirements for hovering from an analysis of the time history of forces and moments in all six degrees of freedom. The wing kinematics necessary to generate sufficient lift are highly constrained by the requirement to balance thrust and pitch torque over the stroke cycle. We also compare the wing motion and aerodynamic forces of free and tethered flies. Tethering causes a strong distortion of the stroke pattern that results in a reduction of translational forces and a prominent nose-down pitch moment. The stereotyped distortion under tethered conditions is most likely due to a disruption of sensory feedback. Finally, we calculated flight power based directly on the measurements of wing motion and aerodynamic forces, which yielded a higher estimate of muscle power during free hovering flight than prior estimates based on time-averaged parameters. This discrepancy is mostly due to a two- to threefold underestimate of the mean profile drag coefficient in prior studies. We also compared our values with the predictions of the same time-averaged models using more accurate kinematic and aerodynamic input parameters based on our high-speed videography measurements. In this case, the time-averaged models tended to overestimate flight
NASA Technical Reports Server (NTRS)
Monta, W. J.
1977-01-01
An experimental investigation was conducted on a model of a wing control version of the Sparrow III type missile to determine the static aerodynamic characteristics over an angle of attack range from 0 deg to 40 deg for Mach numbers from 1.50 to 4.60.
Fluidic Control of Aerodynamic Forces on an Axisymmetric Body
NASA Astrophysics Data System (ADS)
Abramson, Philip; Vukasinovic, Bojan; Glezer, Ari
2007-11-01
The aerodynamic forces and moments on a wind tunnel model of an axisymmetric bluff body are modified by induced local vectoring of the separated base flow. Control is effected by an array of four integrated aft-facing synthetic jets that emanate from narrow, azimuthally-segmented slots, equally distributed around the perimeter of the circular tail end within a small backward facing step that extends into a Coanda surface. The model is suspended in the wind tunnel by eight thin wires for minimal support interference with the wake. Fluidic actuation results in a localized, segmented vectoring of the separated base flow along the rear Coanda surface and induces asymmetric aerodynamic forces and moments to effect maneuvering during flight. The aerodynamic effects associated with quasi-steady and transitory differential, asymmetric activation of the Coanda effect are characterized using direct force and PIV measurements.
Grid sensitivity for aerodynamic optimization and flow analysis
NASA Technical Reports Server (NTRS)
Sadrehaghighi, I.; Tiwari, S. N.
1993-01-01
After reviewing relevant literature, it is apparent that one aspect of aerodynamic sensitivity analysis, namely grid sensitivity, has not been investigated extensively. The grid sensitivity algorithms in most of these studies are based on structural design models. Such models, although sufficient for preliminary or conceptional design, are not acceptable for detailed design analysis. Careless grid sensitivity evaluations, would introduce gradient errors within the sensitivity module, therefore, infecting the overall optimization process. Development of an efficient and reliable grid sensitivity module with special emphasis on aerodynamic applications appear essential. The organization of this study is as follows. The physical and geometric representations of a typical model are derived in chapter 2. The grid generation algorithm and boundary grid distribution are developed in chapter 3. Chapter 4 discusses the theoretical formulation and aerodynamic sensitivity equation. The method of solution is provided in chapter 5. The results are presented and discussed in chapter 6. Finally, some concluding remarks are provided in chapter 7.
Hysteresis zone or locus - Aerodynamic of bulbous based bodies at low speeds
NASA Technical Reports Server (NTRS)
Covert, E. E.
1979-01-01
Experimental data are presented which seem to suggest that a well-defined hysteresis locus on bulbous based bodies at low speeds does not exist. Instead, if the experiment is repeated several times, the entire hysteresis region seems to fill with data rather than trace out a specific hysteresis locus. Data obtained on an oscillating model even at low reduced frequencies may be well defined but when applied to arbitrary motion lead to less accurate results than desired.
NASA Technical Reports Server (NTRS)
Dash, S. M.; York, B. J.; Sinha, N.; Dvorak, F. A.
1987-01-01
An overview of parabolic and PNS (Parabolized Navier-Stokes) methodology developed to treat highly curved sub and supersonic wall jets is presented. The fundamental data base to which these models were applied is discussed in detail. The analysis of strong curvature effects was found to require a semi-elliptic extension of the parabolic modeling to account for turbulent contributions to the normal pressure variations, as well as an extension to the turbulence models utilized, to account for the highly enhanced mixing rates observed in situations with large convex curvature. A noniterative, pressure split procedure is shown to extend parabolic models to account for such normal pressure variations in an efficient manner, requiring minimal additional run time over a standard parabolic approach. A new PNS methodology is presented to solve this problem which extends parabolic methodology via the addition of a characteristic base wave solver. Applications of this approach to analyze the interaction of wave and turbulence processes in wall jets is presented.
Sullivan, W. N.; Leonard, T. M.
1980-11-01
An important aspect of structural design of the Darrieus rotor is the determination of aerodynamic blade loads. This report describes a load generator which has been used at Sandia for quasi-static and dynamic rotor analyses. The generator is based on the single streamtube aerodynamic flow model and is constructed as a FORTRAN IV subroutine to facilitate its use in finite element structural models. Input and output characteristics of the subroutine are described and a complete listing is attached as an appendix.
Aerodynamic optimization of an HSCT configuration using variable-complexity modeling
NASA Technical Reports Server (NTRS)
Hutchison, M. G.; Mason, W. H.; Grossman, B.; Haftka, R. T.
1993-01-01
An approach to aerodynamic configuration optimization is presented for the high-speed civil transport (HSCT). A method to parameterize the wing shape, fuselage shape and nacelle placement is described. Variable-complexity design strategies are used to combine conceptual and preliminary-level design approaches, both to preserve interdisciplinary design influences and to reduce computational expense. Conceptual-design-level (approximate) methods are used to estimate aircraft weight, supersonic wave drag and drag due to lift, and landing angle of attack. The drag due to lift, wave drag and landing angle of attack are also evaluated using more detailed, preliminary-design-level techniques. New, approximate methods for estimating supersonic wave drag and drag due to lift are described. The methodology is applied to the minimization of the gross weight of an HSCT that flies at Mach 2.4 with a range of 5500 n.mi. Results are presented for wing planform shape optimization and for combined wing and fuselage optimization with nacelle placement. Case studies include both all-metal wings and advanced composite wings.
Modeling of space shuttle SRB aft ends for inherent aerodynamic bias determination
NASA Astrophysics Data System (ADS)
González, David R.; Stapf, Sean P.; Gebhard, Thomas J.
2007-04-01
The Air Force's 45th Space Wing is in charge of operating the Range Safety System (RSS) for all launches that take place on the Eastern Range. If initiated, the RSS currently implemented on the Space Transportation System after launch would provide for the partial destruction of the solid rocket boosters (SRBs) to terminate thrust. The majority of the risk from the large explosive debris created comes from the aft ends of the SRBs, which fall largely intact along with the remaining propellant. Historically, no impact data on such a scenario has been available and in support of the Space Shuttle Return-to-Flight schedule, aerodynamic and trajectory analyses were performed to characterize any pitch angle biases associated with the aft end's descent after initiating the linear shaped charges (LSCs) on the SRBs. Results show the aft end has a bias towards impacting at +/-5, 70, or 175 degrees and takes an average of 10 seconds to stabilize into any one of these orientations after being separated from the SRB forward body.
Pickup, B.A.; Thomson, S.L.
2012-01-01
The influence of asymmetric vocal fold stiffness on voice production was evaluated using life-sized, self-oscillating vocal fold models with an idealized geometry based on the human vocal folds. The models were fabricated using flexible, materially-linear silicone compounds with Young’s modulus values comparable to that of vocal fold tissue. The models included a two-layer design to simulate the vocal fold layered structure. The respective Young’s moduli of elasticity of the “left” and “right” vocal fold models were varied to create asymmetric conditions. High-speed videokymography was used to measure maximum vocal fold excursion, vibration frequency, and left-right phase shift, all of which were significantly influenced by asymmetry. Onset pressure, a measure of vocal effort, increased with asymmetry. Particle image velocimetry (PIV) analysis showed significantly greater skewing of the glottal jet in the direction of the stiffer vocal fold model. Potential applications to various clinical conditions are mentioned, and suggestions for future related studies are presented. PMID:19664777
The aerodynamics of small Reynolds numbers
NASA Technical Reports Server (NTRS)
Schmitz, F. W.
1980-01-01
Aerodynamic characteristics of wing model gliders and bird wings in particular are discussed. Wind tunnel measurements and aerodynamics of small Reynolds numbers are enumerated. Airfoil behavior in the critical transition from laminar to turbulent boundary layer, which is more important to bird wing models than to large airplanes, was observed. Experimental results are provided, and an artificial bird wing is described.
NASA Technical Reports Server (NTRS)
Letko, W.
1956-01-01
An experimental investigation has been made in the Langley stability tunnel to determine the aerodynamic characteristics of the Army Chemical Corps model E-112 bomblets with span-chord ratio of 2:1. A detailed analysis has not been made; however, the results showed that all the models were spirally unstable and that a large gap between the model tips and end plates tended to reduce the instability.
X-33 Hypersonic Aerodynamic Characteristics
NASA Technical Reports Server (NTRS)
Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.
1999-01-01
Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will design, build, and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604BOO02G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate the aerodynamic flight database for the hypersonic regime. The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. Al these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.
X-33 Hypersonic Aerodynamic Characteristics
NASA Technical Reports Server (NTRS)
Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.
1999-01-01
Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will build and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604B0002G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate an aerodynamic flight database i n the hypersonic regime. The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. At these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.
X-33 Hypersonic Aerodynamic Characteristics
NASA Technical Reports Server (NTRS)
Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.
1999-01-01
Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will build and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604B0002G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate an aerodynamic flight database in the hypersonic regime. The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. At these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.
X-33 Hypersonic Aerodynamic Characteristics
NASA Technical Reports Server (NTRS)
Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.
1999-01-01
Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will build and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604B0002G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate an aerodynamic flight database in the hypersonic regime, The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. At these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.
NASA Technical Reports Server (NTRS)
Kelley, Henry L.; Crowell, Cynthia A.; Wilson, John C.
1992-01-01
A wind-tunnel investigation was conducted to determine 2-D aerodynamic characteristics of nine polygon-shaped models applicable to helicopter fuselages. The models varied from 1/2 to 1/5 scale and were nominally triangular, diamond, and rectangular in shape. Side force and normal force were obtained at increments of angle of flow incidence from -45 to 90 degrees. The data were compared with results from a baseline UH-60 tail-boom cross-section model. The results indicate that the overall shapes of the plots of normal force and side force were similar to the characteristic shape of the baseline data; however, there were important differences in magnitude. At a flow incidence of 0 degrees, larger values of normal force for the polygon models indicate an increase in fuselage down load of 1 to 2.5 percent of main-rotor thrust compared with the baseline value. Also, potential was indicated among some of the configurations to produce high fuselage side forces and yawing moments compared with the baseline model.
NASA Technical Reports Server (NTRS)
Klein, Vladislav; Noderer, Keith D.
1996-01-01
A nonlinear least squares algorithm for aircraft parameter estimation from flight data was developed. The postulated model for the analysis represented longitudinal, short period motion of an aircraft. The corresponding aerodynamic model equations included indicial functions (unsteady terms) and conventional stability and control derivatives. The indicial functions were modeled as simple exponential functions. The estimation procedure was applied in five examples. Four of the examples used simulated and flight data from small amplitude maneuvers to the F-18 HARV and X-31A aircraft. In the fifth example a rapid, large amplitude maneuver of the X-31 drop model was analyzed. From data analysis of small amplitude maneuvers ft was found that the model with conventional stability and control derivatives was adequate. Also, parameter estimation from a rapid, large amplitude maneuver did not reveal any noticeable presence of unsteady aerodynamics.
NASA Technical Reports Server (NTRS)
Dziubala, T.; Esparza, V.; Gillins, R. L.; Petrozzi, M.
1975-01-01
A Rockwell built 0.030-scale 45-0 modified Space Shuttle Orbiter Configuration 14?A/B model and a Boeing built 0.030-scale 747 carrier model were tested to provide six component force and moment data for each vehicle in proximity to the other at a matrix of relative positions, attitudes and test conditions (angles of attack and sideslip were varied). Orbiter model support system tare effects were determined for corrections to obtain support-free aerodynamics. In addition to the balance force data, pressures were measured. Pressure orifices were located at the base of the Orbiter, on either side of the vertical blade strut, and at the mid-root chord on either side of the vertical tail. Strain gages were installed on the Boeing 747 vertical tail to indicate buffet onset. Photographs of aerodynamic configurations tested are shown.
Ortega, Jason M.; Sabari, Kambiz
2006-03-07
An apparatus for reducing the aerodynamic base drag of a bluff body having a leading end, a trailing end, a top surface, opposing left and right side surfaces, and a base surface at the trailing end substantially normal to a longitudinal centerline of the bluff body, with the base surface joined (1) to the left side surface at a left trailing edge, (2) to the right side surface at a right trailing edge, and (3) to the top surface at a top trailing edge. The apparatus includes left and right vertical boattail plates which are orthogonally attached to the base surface of the bluff body and inwardly offset from the left and right trailing edges, respectively. This produces left and right vertical channels which generate, in a flowstream substantially parallel to the longitudinal centerline, respective left and right vertically-aligned vortical structures, with the left and right vertical boattail plates each having a plate width defined by a rear edge of the plate spaced from the base surface. Each plate also has a peak plate width at a location between top and bottom ends of the plate corresponding to a peak vortex of the respective vertically-aligned vortical structures.
NASA Astrophysics Data System (ADS)
Zeng, Xiao-Hui; Wu, Han; Lai, Jiang; Sheng, Hong-Zhi
2014-12-01
The influences of steady aerodynamic loads on hunting stability of high-speed railway vehicles were investigated in this study. A mechanism is suggested to explain the change of hunting behavior due to actions of aerodynamic loads: the aerodynamic loads can change the position of vehicle system (consequently the contact relations), the wheel/rail normal contact forces, the gravitational restoring forces/moments and the creep forces/moments. A mathematical model for hunting stability incorporating such influences was developed. A computer program capable of incorporating the effects of aerodynamic loads based on the model was written, and the critical speeds were calculated using this program. The dependences of linear and nonlinear critical speeds on suspension parameters considering aerodynamic loads were analyzed by using the orthogonal test method, the results were also compared with the situations without aerodynamic loads. It is shown that the most dominant factors affecting linear and nonlinear critical speeds are different whether the aerodynamic loads considered or not. The damping of yaw damper is the most dominant influencing factor for linear critical speeds, while the damping of lateral damper is most dominant for nonlinear ones. When the influences of aerodynamic loads are considered, the linear critical speeds decrease with the rise of crosswind velocity, whereas it is not the case for the nonlinear critical speeds. The variation trends of critical speeds with suspension parameters can be significantly changed by aerodynamic loads. Combined actions of aerodynamic loads and suspension parameters also affect the critical speeds. The effects of such joint action are more obvious for nonlinear critical speeds.
Flow Induced Vibration and Glottal Aerodynamics in a Three-Dimensional Laryngeal Model
NASA Astrophysics Data System (ADS)
Zheng, Xudong; Xue, Qian; Mittal, Rajat; Bielamowicz, Steven
2009-11-01
Three-dimensional effects associated with phonation remain unclear due to the lack of capability of simulating 3D fluid-tissue interaction in the past. To advance the state-of-the-art in this arena, an immersed-boundary method based flow solver coupled with a finite-element solid dynamics solver is employed to conduct high-fidelity direct-numerical simulations of phonation in a 3D model of the human larynx. Three-dimensional vibration patterns are captured along with turbulence effects and three-dimensional vortex structures in the glottal jet. Results from these simulations are presented.
NASA Technical Reports Server (NTRS)
Dwoyer, D. L.; Newman, P. A.; Thames, F. C.; Melson, N. D.
1981-01-01
The problem of predicting aerodynamic loads on the insulating tiles of the space shuttle thermal protection system (TPS) is discussed and seen to require a method for predicting pressure and mass flux in the gaps between tiles. A mathematical model of the tile-gap flow is developed, based upon a slow viscous (Stokes) flow analysis, and is verified against experimental data. The tile-gap pressure field is derived from a solution of the two-dimensional Laplace equation; the mass-flux vector is then calculated from the pressure gradient. The means for incorporating this model into a lumped-parameter network analogy for porous-media flow is given. The means for incorporating this model into a lumped-parameter network analogy for porous-media flow is given. The flow model shows tile-gap mass flux to be very sensitive to the gap width indicating a need for coupling the TPS flow and tile displacement calculation. Analytical and experimental work to improve TPS flow predictions and a possible shuttle TPS hardware modification are recommended.
NASA Astrophysics Data System (ADS)
Suzuki, Kensuke
A new analysis tool, an unsteady Hybrid Navier-Stokes/Vortex Model, for a horizontal axis wind turbine (HAWT) in yawed flow is presented, and its convergence and low cost computational performance are demonstrated. In earlier work, a steady Hybrid Navier-Stokes/Vortex Model was developed with a view to improving simulation results obtained by participants of the NASA Ames blind comparison workshop, following the NREL Unsteady Aerodynamics Experiment. The hybrid method was shown to better predict rotor torque and power over the range of wind speeds, from fully attached to separated flows. A decade has passed since the workshop was held and three dimensional unsteady Navier-Stokes analyses have become available using super computers. In the first chapter, recent results of unsteady Euler and Navier-Stokes computations are reviewed as standard references of what is currently possible and are contrasted with results of the Hybrid Navier-Stokes/Vortex Model in steady flow. In Chapter 2, the computational method for the unsteady Hybrid model is detailed. The grid generation procedure, using ICEM CFD, is presented in Chapter 3. Steady and unsteady analysis results for the NREL Phase IV rotor and for a modified "swept NREL rotor" are presented in Chapter 4-Chapter 7.
Aerodynamic Simulation of Runback Ice Accretion
NASA Technical Reports Server (NTRS)
Broeren, Andy P.; Whalen, Edward A.; Busch, Greg T.; Bragg, Michael B.
2010-01-01
This report presents the results of recent investigations into the aerodynamics of simulated runback ice accretion on airfoils. Aerodynamic tests were performed on a full-scale model using a high-fidelity, ice-casting simulation at near-flight Reynolds (Re) number. The ice-casting simulation was attached to the leading edge of a 72-in. (1828.8-mm ) chord NACA 23012 airfoil model. Aerodynamic performance tests were conducted at the ONERA F1 pressurized wind tunnel over a Reynolds number range of 4.7?10(exp 6) to 16.0?10(exp 6) and a Mach (M) number ran ge of 0.10 to 0.28. For Re = 16.0?10(exp 6) and M = 0.20, the simulated runback ice accretion on the airfoil decreased the maximum lift coe fficient from 1.82 to 1.51 and decreased the stalling angle of attack from 18.1deg to 15.0deg. The pitching-moment slope was also increased and the drag coefficient was increased by more than a factor of two. In general, the performance effects were insensitive to Reynolds numb er and Mach number changes over the range tested. Follow-on, subscale aerodynamic tests were conducted on a quarter-scale NACA 23012 model (18-in. (457.2-mm) chord) at Re = 1.8?10(exp 6) and M = 0.18, using low-fidelity, geometrically scaled simulations of the full-scale castin g. It was found that simple, two-dimensional simulations of the upper- and lower-surface runback ridges provided the best representation of the full-scale, high Reynolds number iced-airfoil aerodynamics, whereas higher-fidelity simulations resulted in larger performance degrada tions. The experimental results were used to define a new subclassification of spanwise ridge ice that distinguishes between short and tall ridges. This subclassification is based upon the flow field and resulting aerodynamic characteristics, regardless of the physical size of the ridge and the ice-accretion mechanism.
Recent Enhancements to the Development of CFD-Based Aeroelastic Reduced-Order Models
NASA Technical Reports Server (NTRS)
Silva, Walter A.
2007-01-01
Recent enhancements to the development of CFD-based unsteady aerodynamic and aeroelastic reduced-order models (ROMs) are presented. These enhancements include the simultaneous application of structural modes as CFD input, static aeroelastic analysis using a ROM, and matched-point solutions using a ROM. The simultaneous application of structural modes as CFD input enables the computation of the unsteady aerodynamic state-space matrices with a single CFD execution, independent of the number of structural modes. The responses obtained from a simultaneous excitation of the CFD-based unsteady aerodynamic system are processed using system identification techniques in order to generate an unsteady aerodynamic state-space ROM. Once the unsteady aerodynamic state-space ROM is generated, a method for computing the static aeroelastic response using this unsteady aerodynamic ROM and a state-space model of the structure, is presented. Finally, a method is presented that enables the computation of matchedpoint solutions using a single ROM that is applicable over a range of dynamic pressures and velocities for a given Mach number. These enhancements represent a significant advancement of unsteady aerodynamic and aeroelastic ROM technology.
Contribution to the aerodynamic study of wings and propellers
NASA Technical Reports Server (NTRS)
Menard, M.
1983-01-01
Various problems regarding the aerodynamics of lifting wings are solved. Two methods are proposed for replacing the wing, both involving "viscous" edge vortices. the applications give results which agree well with experiments. Two new methods are also proposed for calculating propellers based on the vortex model consisting of an edge vortex and a "viscous" hub vortex.
NASA Technical Reports Server (NTRS)
Waszak, Martin R.; Fung, Jimmy
1998-01-01
This report describes the development of transfer function models for the trailing-edge and upper and lower spoiler actuators of the Benchmark Active Control Technology (BACT) wind tunnel model for application to control system analysis and design. A simple nonlinear least-squares parameter estimation approach is applied to determine transfer function parameters from frequency response data. Unconstrained quasi-Newton minimization of weighted frequency response error was employed to estimate the transfer function parameters. An analysis of the behavior of the actuators over time to assess the effects of wear and aerodynamic load by using the transfer function models is also presented. The frequency responses indicate consistent actuator behavior throughout the wind tunnel test and only slight degradation in effectiveness due to aerodynamic hinge loading. The resulting actuator models have been used in design, analysis, and simulation of controllers for the BACT to successfully suppress flutter over a wide range of conditions.
Improved Aerodynamic Influence Coefficients for Dynamic Aeroelastic Analyses
NASA Astrophysics Data System (ADS)
Gratton, Patrice
2011-12-01
Currently at Bombardier Aerospace, aeroelastic analyses are performed using the Doublet Lattice Method (DLM) incorporated in the NASTRAN solver. This method proves to be very reliable and fast in preliminary design stages where wind tunnel experimental results are often not available. Unfortunately, the geometric simplifications and limitations of the DLM, based on the lifting surfaces theory, reduce the ability of this method to give reliable results for all flow conditions, particularly in transonic flow. Therefore, a new method has been developed involving aerodynamic data from high-fidelity CFD codes which solve the Euler or Navier-Stokes equations. These new aerodynamic loads are transmitted to the NASTRAN aeroelastic module through improved aerodynamic influence coefficients (AIC). A cantilevered wing model is created from the Global Express structural model and a set of natural modes is calculated for a baseline configuration of the structure. The baseline mode shapes are then combined with an interpolation scheme to deform the 3-D CFD mesh necessary for Euler and Navier-Stokes analyses. An uncoupled approach is preferred to allow aerodynamic information from different CFD codes. Following the steady state CFD analyses, pressure differences ( DeltaCp), calculated between the deformed models and the original geometry, lead to aerodynamic loads which are transferred to the DLM model. A modal-based AIC method is applied to the aerodynamic matrices of NASTRAN based on a least-square approximation to evaluate aerodynamic loads of a different wing configuration which displays similar types of mode shapes. The methodology developed in this research creates weighting factors based on steady CFD analyses which have an equivalent reduced frequency of zero. These factors are applied to both the real and imaginary part of the aerodynamic matrices as well as all reduced frequencies used in the PK-Method which solves flutter problems. The modal-based AIC method
Supersonic Aerodynamic Characteristics of Proposed Mars '07 Smart Lander Configurations
NASA Technical Reports Server (NTRS)
Murphy, Kelly J.; Horvath, Thomas J.; Erickson, Gary E.; Green, Joseph M.
2002-01-01
Supersonic aerodynamic data were obtained for proposed Mars '07 Smart Lander configurations in NASA Langley Research Center's Unitary Plan Wind Tunnel. The primary objective of this test program was to assess the supersonic aerodynamic characteristics of the baseline Smart Lander configuration with and without fixed shelf/tab control surfaces. Data were obtained over a Mach number range of 2.3 to 4.5, at a free stream Reynolds Number of 1 x 10(exp 6) based on body diameter. All configurations were run at angles of attack from -5 to 20 degrees and angles of sideslip of -5 to 5 degrees. These results were complemented with computational fluid dynamic (CFD) predictions to enhance the understanding of experimentally observed aerodynamic trends. Inviscid and viscous full model CFD solutions compared well with experimental results for the baseline and 3 shelf/tab configurations. Over the range tested, Mach number effects were shown to be small on vehicle aerodynamic characteristics. Based on the results from 3 different shelf/tab configurations, a fixed control surface appears to be a feasible concept for meeting aerodynamic performance metrics necessary to satisfy mission requirements.
Modeling the transient aerodynamic effects during the motion of a flexible trailing edge
NASA Astrophysics Data System (ADS)
Wolff, T.; Seume, J. R.
2016-09-01
Wind turbine blades have been becoming longer and more slender during the last few decades. The longer lever arm results in higher stresses at the blade root. Hence, the unsteady loads induced by turbulence, gust, or wind shear increase. One promising way to control these loads is to use flexible trailing edges near the blade tip. The unsteady effects which appear during the motion of a flexible trailing edge must be considered for the load calculation during the design process because of their high influence on aeroelastic effects and hence on the fatigue loads. This is not yet possible in most of the wind turbine simulation environments. Consequently, an empirical model is developed in the present study which accounts for unsteady effects during the motion of the trailing edge. The model is based on Fourier analyses of results generated with Reynolds-Averaged Navier-Stokes (RANS) simulations of a typical thin airfoil with a deformable trailing edge. The validation showed that the model fits Computational Fluid Dynamics (CFD) results simulated with a random time series of the deflection angle.
NASA Technical Reports Server (NTRS)
Blair, A. B., Jr.
1972-01-01
An investigation has been made in the Langley Unitary Plan wind tunnel to determine the aerodynamic characteristics of a lifting-body orbiter model with a blunted delta planform. The model was tested at Mach numbers from 2.30 to 4.60, at nominal angles of attack from -4 deg to 60 deg and angles of sideslip from -4 deg to 10 deg, and at a Reynolds number of 2.5 million per foot.
NASA Technical Reports Server (NTRS)
Ganzer, Victor M
1944-01-01
Results are presented for tests of two wings, an NACA 230-series wing and a highly-cambered NACA 66-series wing on a twin-engine pursuit airplane. Auxiliary control flaps were tested in combinations with each wing. Data showing comparison of high-speed aerodynamic characteristics of the model when equipped with each wing, the effect of the auxiliary control flaps on aerodynamic characteristics, and elevator effectiveness for the model with the 66-series wing are presented. High-speed aerodynamic characteristics of the model were improved with the 66-series wing.
Motion transitions of falling plates via quasisteady aerodynamics.
Hu, Ruifeng; Wang, Lifeng
2014-07-01
In this paper, we study the dynamics of freely falling plates based on the Kirchhoff equation and the quasisteady aerodynamic model. Motion transitions among fluttering, tumbling along a cusp-like trajectory, irregular, and tumbling along a straight trajectory are obtained by solving the dynamical equations. Phase diagrams spanning between the nondimensional moment of inertia and aerodynamic coefficients or aspect ratio are built to identify regimes for these falling styles. We also investigate the stability of fixed points and bifurcation scenarios. It is found that the transitions are all heteroclinic bifurcations and the influence of the fixed-point stability is local.
Aeroacoustics. [analysis of properties of sound generated by aerodynamic forces
NASA Technical Reports Server (NTRS)
Goldstein, M., E.
1974-01-01
An analysis was conducted to determine the properties of sound generated by aerodynamic forces or motions originating in a flow, such as the unsteady aerodynamic forces on propellers or by turbulent flows around an aircraft. The acoustics of moving media are reviewed and mathematical models are developed. Lighthill's acoustic analogy and the application to turbulent flows are analyzed. The effects of solid boundaries are calculated. Theories based on the solution of linearized vorticity and acoustic field equations are explained. The effects of nonuniform mean flow on the generation of sound are reported.
NASA Technical Reports Server (NTRS)
Erickson, Gary E.; Inenaga, Andrew S.
1994-01-01
Laser vapor screen (LVS) flow visualization systems that are fiber-optic based were developed and installed for aerodynamic research in the Langley 8-Foot Transonic Pressure Tunnel and the Langley 7- by 10-Foot High Speed Tunnel. Fiber optics are used to deliver the laser beam through the plenum shell that surrounds the test section of each facility and to the light-sheet-generating optics positioned in the ceiling window of the test section. Water is injected into the wind tunnel diffuser section to increase the relative humidity and promote condensation of the water vapor in the flow field about the model. The condensed water vapor is then illuminated with an intense sheet of laser light to reveal features of the flow field. The plenum shells are optically sealed; therefore, video-based systems are used to observe and document the flow field. Operational experience shows that the fiber-optic-based systems provide safe, reliable, and high-quality off-surface flow visualization in smaller and larger scale subsonic and transonic wind tunnels. The design, the installation, and the application of the Langley Research Center (LaRC) LVS flow visualization systems in larger scale wind tunnels are highlighted. The efficiency of the fiber optic LVS systems and their insensitivity to wind tunnel vibration, the tunnel operating temperature and pressure variations, and the airborne contaminants are discussed.
NASA Technical Reports Server (NTRS)
Jones, R. T. (Compiler)
1979-01-01
A collection of papers on modern theoretical aerodynamics is presented. Included are theories of incompressible potential flow and research on the aerodynamic forces on wing and wing sections of aircraft and on airship hulls.
NASA Technical Reports Server (NTRS)
Vicker, Darby
2006-01-01
A viewgraph presentation describing aerodynamics at NASA Johnson Space Center is shown. The topics include: 1) Personal Background; 2) Aerodynamic Tools; 3) The Overset Computational Fluid Dynamics (CFD) Process; and 4) Recent Applicatoins.
NASA Technical Reports Server (NTRS)
Kral, Linda D.; Ladd, John A.; Mani, Mori
1995-01-01
The objective of this viewgraph presentation is to evaluate turbulence models for integrated aircraft components such as the forebody, wing, inlet, diffuser, nozzle, and afterbody. The one-equation models have replaced the algebraic models as the baseline turbulence models. The Spalart-Allmaras one-equation model consistently performs better than the Baldwin-Barth model, particularly in the log-layer and free shear layers. Also, the Sparlart-Allmaras model is not grid dependent like the Baldwin-Barth model. No general turbulence model exists for all engineering applications. The Spalart-Allmaras one-equation model and the Chien k-epsilon models are the preferred turbulence models. Although the two-equation models often better predict the flow field, they may take from two to five times the CPU time. Future directions are in further benchmarking the Menter blended k-w/k-epsilon and algorithmic improvements to reduce CPU time of the two-equation model.
NASA Iced Aerodynamics and Controls Current Research
NASA Technical Reports Server (NTRS)
Addy, Gene
2009-01-01
This slide presentation reviews the state of current research in the area of aerodynamics and aircraft control with ice conditions by the Aviation Safety Program, part of the Integrated Resilient Aircraft Controls Project (IRAC). Included in the presentation is a overview of the modeling efforts. The objective of the modeling is to develop experimental and computational methods to model and predict aircraft response during adverse flight conditions, including icing. The Aircraft icing modeling efforts includes the Ice-Contaminated Aerodynamics Modeling, which examines the effects of ice contamination on aircraft aerodynamics, and CFD modeling of ice-contaminated aircraft aerodynamics, and Advanced Ice Accretion Process Modeling which examines the physics of ice accretion, and works on computational modeling of ice accretions. The IRAC testbed, a Generic Transport Model (GTM) and its use in the investigation of the effects of icing on its aerodynamics is also reviewed. This has led to a more thorough understanding and models, both theoretical and empirical of icing physics and ice accretion for airframes, advanced 3D ice accretion prediction codes, CFD methods for iced aerodynamics and better understanding of aircraft iced aerodynamics and its effects on control surface effectiveness.
NASA Technical Reports Server (NTRS)
Williams, Louis J.; Hessenius, Kristin A.; Corsiglia, Victor R.; Hicks, Gary; Richardson, Pamela F.; Unger, George; Neumann, Benjamin; Moss, Jim
1992-01-01
The annual accomplishments is reviewed for the Aerodynamics Division during FY 1991. The program includes both fundamental and applied research directed at the full spectrum of aerospace vehicles, from rotorcraft to planetary entry probes. A comprehensive review is presented of the following aerodynamics elements: computational methods and applications; CFD validation; transition and turbulence physics; numerical aerodynamic simulation; test techniques and instrumentation; configuration aerodynamics; aeroacoustics; aerothermodynamics; hypersonics; subsonics; fighter/attack aircraft and rotorcraft.
NASA Technical Reports Server (NTRS)
Holmes, Bruce J.; Schairer, Edward; Hicks, Gary; Wander, Stephen; Blankson, Isiaiah; Rose, Raymond; Olson, Lawrence; Unger, George
1990-01-01
Presented here is a comprehensive review of the following aerodynamics elements: computational methods and applications, computational fluid dynamics (CFD) validation, transition and turbulence physics, numerical aerodynamic simulation, drag reduction, test techniques and instrumentation, configuration aerodynamics, aeroacoustics, aerothermodynamics, hypersonics, subsonic transport/commuter aviation, fighter/attack aircraft and rotorcraft.
Aerodynamic design of electric and hybrid vehicles: A guidebook
NASA Technical Reports Server (NTRS)
Kurtz, D. W.
1980-01-01
A typical present-day subcompact electric hybrid vehicle (EHV), operating on an SAE J227a D driving cycle, consumes up to 35% of its road energy requirement overcoming aerodynamic resistance. The application of an integrated system design approach, where drag reduction is an important design parameter, can increase the cycle range by more than 15%. This guidebook highlights a logic strategy for including aerodynamic drag reduction in the design of electric and hybrid vehicles to the degree appropriate to the mission requirements. Backup information and procedures are included in order to implement the strategy. Elements of the procedure are based on extensive wind tunnel tests involving generic subscale models and full-scale prototype EHVs. The user need not have any previous aerodynamic background. By necessity, the procedure utilizes many generic approximations and assumptions resulting in various levels of uncertainty. Dealing with these uncertainties, however, is a key feature of the strategy.
Aerodynamic drag on intermodal railcars
NASA Astrophysics Data System (ADS)
Kinghorn, Philip; Maynes, Daniel
2014-11-01
The aerodynamic drag associated with transport of commodities by rail is becoming increasingly important as the cost of diesel fuel increases. This study aims to increase the efficiency of intermodal cargo trains by reducing the aerodynamic drag on the load carrying cars. For intermodal railcars a significant amount of aerodynamic drag is a result of the large distance between loads that often occurs and the resulting pressure drag resulting from the separated flow. In the present study aerodynamic drag data have been obtained through wind tunnel testing on 1/29 scale models to understand the savings that may be realized by judicious modification to the size of the intermodal containers. The experiments were performed in the BYU low speed wind tunnel and the test track utilizes two leading locomotives followed by a set of five articulated well cars with double stacked containers. The drag on a representative mid-train car is measured using an isolated load cell balance and the wind tunnel speed is varied from 20 to 100 mph. We characterize the effect that the gap distance between the containers and the container size has on the aerodynamic drag of this representative rail car and investigate methods to reduce the gap distance.
Single dielectric barrier discharge plasma enhanced aerodynamics: physics, modeling and applications
NASA Astrophysics Data System (ADS)
Corke, Thomas C.; Post, Martiqua L.; Orlov, Dmitriy M.
2009-01-01
The term “plasma actuator” has been a part of the fluid dynamics flow control vernacular for more than a decade. A particular type of plasma actuator that has gained wide use is based on a single dielectric barrier discharge (SDBD) mechanism that has desirable features for use in air at atmospheric pressures. For these actuators, the mechanism of flow control is through a generated body force vector that couples with the momentum in the external flow. The body force can be derived from first principles and the plasma actuator effect can be easily incorporated into flow solvers so that their placement and operation can be optimized. They have been used in a wide range of applications that include bluff body wake control; lift augmentation and separation control on a variety of lifting surfaces ranging from fixed wings with various degrees of sweep, wind turbine rotors and pitching airfoils simulating helicopter rotors; flow separation and tip-casing clearance flow control to reduce losses in turbines, to control flow surge and stall in compressors; and in exciting instabilities in boundary layers at subsonic to supersonic Mach numbers for turbulent transition control. New applications continue to appear through programs in a growing number of US universities and government laboratories, as well as in Germany, France, England, Netherland, Russia, Japan and China. This paper provides an overview of the physics, design and modeling of SDBD plasma actuators. It then presents their use in a number of applications that includes both numerical flow simulations and experiments together.
NASA Technical Reports Server (NTRS)
Capone, Francis J.; Carson, George T., Jr.
1985-01-01
An investigation has been conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of empennage surface location and vertical tail cant angle on the aft-end aerodynamic characteristics of a twin-engine fighter-type configuration. The configuration featured two-dimensional convergent-divergent nozzles and twin-vertical tails. The investigation was conducted with different empennage locations that included two horizontal and three vertical tail positions. Vertical tail cant angle was varied from -10 deg to 20 deg for one selected configuration. Tests were conducted at Mach number 0.60 to 1.20 and at angles of attack -3 to 9 deg. Nozzle pressure ratio was varied from jet off to approximately 9, depending upon Mach number. Tail interference effects were present throughout the range of Mach numbers tested and found to be either favorable or adverse, depending upon test condition and model configuration. At a Mach number of 0.90, adverse interference effects accounted for a significant percentage of total aft-end drag. Interference effects on the nozzle were generally favorable but became adverse as the horizontal tails were moved from a mid to an aft position. The configuration with nonaxisymmetric nozzles had lower total aft-end drag with tails-off than a similar configuration with axisymmetric nozzles at Mach numbers of 0.60 and 0.90.
NASA Astrophysics Data System (ADS)
Abaimov, N. A.; Ryzhkov, A. F.
2015-11-01
Problems requiring solution in development of modern highly efficient gasification reactor of a promising high power integrated gasification combined-cycle plant are formulated. The task of creating and testing a numerical model of an entrained-flow reactor for thermochemical conversion of pulverized coal is solved. The basic method of investigation is computational fluid dynamics. The submodel of thermochemical processes, including a single-stage scheme of volatile substances outlet and three heterogeneous reactions of carbon residue conversion (complete carbon oxidation, Boudouard reaction and hydrogasification), is given. The mass loss rate is determined according to the basic assumptions of the diffusion-kinetic theory. The equations applied for calculation of the process of outlet of volatile substances and three stages of fuel gasifi-cation (diffusion of reagent gas toward the surface of the coal particle, heterogeneous reactions of gas with carbon on its surface, and homogeneous reactions beyond the particle surface) are presented. The universal combined submodel Eddy Dissipation/Finite Rate Chemistry with standard (built-in) constants is used for numerical estimates. Aerodynamic mechanisms of action on thermochemical processes of solid fuel gasification are studied, as exemplified by the design upgrade of a cyclone reactor of preliminary thermal fuel preparation. Volume concentrations of combustible gases and products of complete combustion in the syngas before and after primary air and pulverized coal flows' redistribution are given. Volume concentrations of CO in syngas at different positions of tangential secondary air inlet nozzle are compared.
NASA Technical Reports Server (NTRS)
Graves, E. B.
1972-01-01
A study has been made to determine the aerodynamic characteristics of a low-aspect ratio cruciform missile model with all-movable wings and tails. The configuration was tested at Mach numbers from 1.50 to 4.63 with the wings in the vertical and horizontal planes and with the wings in a 45 deg roll plane with tails in line and interdigitated.
Aerodynamic Simulation of the MARINTEK Braceless Semisubmersible Wave Tank Tests
NASA Astrophysics Data System (ADS)
Stewart, Gordon; Muskulus, Michael
2016-09-01
Model scale experiments of floating offshore wind turbines are important for both platform design for the industry as well as numerical model validation for the research community. An important consideration in the wave tank testing of offshore wind turbines are scaling effects, especially the tension between accurate scaling of both hydrodynamic and aerodynamic forces. The recent MARINTEK braceless semisubmersible wave tank experiment utilizes a novel aerodynamic force actuator to decouple the scaling of the aerodynamic forces. This actuator consists of an array of motors that pull on cables to provide aerodynamic forces that are calculated by a blade-element momentum code in real time as the experiment is conducted. This type of system has the advantage of supplying realistically scaled aerodynamic forces that include dynamic forces from platform motion, but does not provide the insights into the accuracy of the aerodynamic models that an actual model-scale rotor could provide. The modeling of this system presents an interesting challenge, as there are two ways to simulate the aerodynamics; either by using the turbulent wind fields as inputs to the aerodynamic model of the design code, or by surpassing the aerodynamic model and using the forces applied to the experimental turbine as direct inputs to the simulation. This paper investigates the best practices of modeling this type of novel aerodynamic actuator using a modified wind turbine simulation tool, and demonstrates that bypassing the dynamic aerodynamics solver of design codes can lead to erroneous results.
Free Wake Techniques for Rotor Aerodynamic Analylis. Volume 2: Vortex Sheet Models
NASA Technical Reports Server (NTRS)
Tanuwidjaja, A.
1982-01-01
Results of computations are presented using vortex sheets to model the wake and test the sensitivity of the solutions to various assumptions used in the development of the models. The complete codings are included.
Electrical and kinetic model of an atmospheric rf device for plasma aerodynamics applications
Pinheiro, Mario J.; Martins, Alexandre A.
2010-08-15
The asymmetrically mounted flat plasma actuator is investigated using a self-consistent two-dimensional fluid model at atmospheric pressure. The computational model assumes the drift-diffusion approximation and uses a simple plasma kinetic model. It investigated the electrical and kinetic properties of the plasma, calculated the charged species concentrations, surface charge density, electrohydrodynamic forces, and gas speed. The present computational model contributes to understand the main physical mechanisms, and suggests ways to improve its performance.
Asymmetric Uncertainty Expression for High Gradient Aerodynamics
NASA Technical Reports Server (NTRS)
Pinier, Jeremy T
2012-01-01
When the physics of the flow around an aircraft changes very abruptly either in time or space (e.g., flow separation/reattachment, boundary layer transition, unsteadiness, shocks, etc), the measurements that are performed in a simulated environment like a wind tunnel test or a computational simulation will most likely incorrectly predict the exact location of where (or when) the change in physics happens. There are many reasons for this, includ- ing the error introduced by simulating a real system at a smaller scale and at non-ideal conditions, or the error due to turbulence models in a computational simulation. The un- certainty analysis principles that have been developed and are being implemented today do not fully account for uncertainty in the knowledge of the location of abrupt physics changes or sharp gradients, leading to a potentially underestimated uncertainty in those areas. To address this problem, a new asymmetric aerodynamic uncertainty expression containing an extra term to account for a phase-uncertainty, the magnitude of which is emphasized in the high-gradient aerodynamic regions is proposed in this paper. Additionally, based on previous work, a method for dispersing aerodynamic data within asymmetric uncer- tainty bounds in a more realistic way has been developed for use within Monte Carlo-type analyses.
Aerodynamic detuning analysis of an unstalled supersonic turbofan cascade
NASA Technical Reports Server (NTRS)
Hoyniak, D.; Fleeter, S.
1985-01-01
An approach to passive flutter control is aerodynamic detuning, defined as designed passage-to-passage differences in the unsteady aerodynamic flow field of a rotor blade row. Thus, aerodynamic detuning directly affects the fundamental driving mechanism for flutter. A model to demonstrate the enhanced supersonic aeroelastic stability associated with aerodynamic detuning is developed. The stability of an aerodynamically detuned cascade operating in a supersonic inlet flow field with a subsonic leading edge locus is analyzed, with the aerodynamic detuning accomplished by means of nonuniform circumferential spacing of adjacent rotor blades. The unsteady aerodynamic forces and moments on the blading are defined in terms of influence coefficients in a manner that permits the stability of both a conventional uniformally spaced rotor configuration as well as the detuned nonuniform circumferentially spaced rotor to be determined. With Verdon's uniformly spaced Cascade B as a baseline, this analysis is then utilized to demonstrate the potential enhanced aeroelastic stability associated with this particular type of aerodynamic detuning.
Aerodynamics Research Revolutionizes Truck Design
NASA Technical Reports Server (NTRS)
2008-01-01
During the 1970s and 1980s, researchers at Dryden Flight Research Center conducted numerous tests to refine the shape of trucks to reduce aerodynamic drag and improved efficiency. During the 1980s and 1990s, a team based at Langley Research Center explored controlling drag and the flow of air around a moving body. Aeroserve Technologies Ltd., of Ottawa, Canada, with its subsidiary, Airtab LLC, in Loveland, Colorado, applied the research from Dryden and Langley to the development of the Airtab vortex generator. Airtabs create two counter-rotating vortices to reduce wind resistance and aerodynamic drag of trucks, trailers, recreational vehicles, and many other vehicles.
NASA Technical Reports Server (NTRS)
Zahm, A F
1924-01-01
This report gives the description and the use of a specially designed aerodynamic plane table. For the accurate and expeditious geometrical measurement of models in an aerodynamic laboratory, and for miscellaneous truing operations, there is frequent need for a specially equipped plan table. For example, one may have to measure truly to 0.001 inch the offsets of an airfoil at many parts of its surface. Or the offsets of a strut, airship hull, or other carefully formed figure may require exact calipering. Again, a complete airplane model may have to be adjusted for correct incidence at all parts of its surfaces or verified in those parts for conformance to specifications. Such work, if but occasional, may be done on a planing or milling machine; but if frequent, justifies the provision of a special table. For this reason it was found desirable in 1918 to make the table described in this report and to equip it with such gauges and measures as the work should require.
NASA Technical Reports Server (NTRS)
Grossman, Bernard
1999-01-01
The technical details are summarized below: Compressible and incompressible versions of a three-dimensional unstructured mesh Reynolds-averaged Navier-Stokes flow solver have been differentiated and resulting derivatives have been verified by comparisons with finite differences and a complex-variable approach. In this implementation, the turbulence model is fully coupled with the flow equations in order to achieve this consistency. The accuracy demonstrated in the current work represents the first time that such an approach has been successfully implemented. The accuracy of a number of simplifying approximations to the linearizations of the residual have been examined. A first-order approximation to the dependent variables in both the adjoint and design equations has been investigated. The effects of a "frozen" eddy viscosity and the ramifications of neglecting some mesh sensitivity terms were also examined. It has been found that none of the approximations yielded derivatives of acceptable accuracy and were often of incorrect sign. However, numerical experiments indicate that an incomplete convergence of the adjoint system often yield sufficiently accurate derivatives, thereby significantly lowering the time required for computing sensitivity information. The convergence rate of the adjoint solver relative to the flow solver has been examined. Inviscid adjoint solutions typically require one to four times the cost of a flow solution, while for turbulent adjoint computations, this ratio can reach as high as eight to ten. Numerical experiments have shown that the adjoint solver can stall before converging the solution to machine accuracy, particularly for viscous cases. A possible remedy for this phenomenon would be to include the complete higher-order linearization in the preconditioning step, or to employ a simple form of mesh sequencing to obtain better approximations to the solution through the use of coarser meshes. . An efficient surface parameterization based
NASA Technical Reports Server (NTRS)
Grossman, Bernard
1999-01-01
Compressible and incompressible versions of a three-dimensional unstructured mesh Reynolds-averaged Navier-Stokes flow solver have been differentiated and resulting derivatives have been verified by comparisons with finite differences and a complex-variable approach. In this implementation, the turbulence model is fully coupled with the flow equations in order to achieve this consistency. The accuracy demonstrated in the current work represents the first time that such an approach has been successfully implemented. The accuracy of a number of simplifying approximations to the linearizations of the residual have been examined. A first-order approximation to the dependent variables in both the adjoint and design equations has been investigated. The effects of a "frozen" eddy viscosity and the ramifications of neglecting some mesh sensitivity terms were also examined. It has been found that none of the approximations yielded derivatives of acceptable accuracy and were often of incorrect sign. However, numerical experiments indicate that an incomplete convergence of the adjoint system often yield sufficiently accurate derivatives, thereby significantly lowering the time required for computing sensitivity information. The convergence rate of the adjoint solver relative to the flow solver has been examined. Inviscid adjoint solutions typically require one to four times the cost of a flow solution, while for turbulent adjoint computations, this ratio can reach as high as eight to ten. Numerical experiments have shown that the adjoint solver can stall before converging the solution to machine accuracy, particularly for viscous cases. A possible remedy for this phenomenon would be to include the complete higher-order linearization in the preconditioning step, or to employ a simple form of mesh sequencing to obtain better approximations to the solution through the use of coarser meshes. An efficient surface parameterization based on a free-form deformation technique has been
Feedback Control for Aerodynamics (Preprint)
2006-09-01
AFRL-VA-WP-TP-2006-348 FEEDBACK CONTROL FOR AERODYNAMICS (PREPRINT) R. Chris Camphouse, Seddik M. Djouadi, and James H. Myatt...CONSTRUCTION FOR THE DESIGN OF BOUNDARY FEEDBACK CONTROLS FROM REDUCED ORDER MODELS (PREPRINT) 5c. PROGRAM ELEMENT NUMBER 0601102F 5d. PROJECT NUMBER...
Turbulence Model Behavior in Low Reynolds Number Regions of Aerodynamic Flowfields
NASA Technical Reports Server (NTRS)
Rumsey, Christopher L.; Spalart, Philippe R.
2008-01-01
The behaviors of the widely-used Spalart-Allmaras (SA) and Menter shear-stress transport (SST) turbulence models at low Reynolds numbers and under conditions conducive to relaminarization are documented. The flows used in the investigation include 2-D zero pressure gradient flow over a flat plate from subsonic to hypersonic Mach numbers, 2-D airfoil flow from subsonic to supersonic Mach numbers, 2-D subsonic sink-flow, and 3-D subsonic flow over an infinite swept wing (particularly its leading-edge region). Both models exhibit a range over which they behave transitionally in the sense that the flow is neither laminar nor fully turbulent, but these behaviors are different: the SST model typically has a well-defined transition location, whereas the SA model does not. Both models are predisposed to delayed activation of turbulence with increasing freestream Mach number. Also, both models can be made to achieve earlier activation of turbulence by increasing their freestream levels, but too high a level can disturb the turbulent solution behavior. The technique of maintaining freestream levels of turbulence without decay in the SST model, introduced elsewhere, is shown here to be useful in reducing grid-dependence of the model's transitional behavior. Both models are demonstrated to be incapable of predicting relaminarization; eddy viscosities remain weakly turbulent in accelerating or laterally-strained boundary layers for which experiment and direct simulations indicate turbulence suppression. The main conclusion is that these models are intended for fully turbulent high Reynolds number computations, and using them for transitional (e.g., low Reynolds number) or relaminarizing flows is not appropriate.
Turbulence Model Behavior in Low Reynolds Number Regions of Aerodynamic Flowfields
NASA Technical Reports Server (NTRS)
Rumsey, Christopher L.; Spalart, Philippe R.
2008-01-01
The behaviors of the widely-used Spalart-Allmaras (SA) and Menter shear-stress transport (SST) turbulence models at low Reynolds numbers and under conditions conducive to relaminarization are documented. The flows used in the investigation include 2-D zero pressure gradient flow over a flat plate from subsonic to hypersonic Mach numbers, 2-D airfoil flow from subsonic to supersonic Mach numbers, 2-D subsonic sink-flow, and 3-D subsonic flow over an infinite swept wing (particularly its leading-edge region). Both models exhibit a range over which they behave 'transitionally' in the sense that the flow is neither laminar nor fully turbulent, but these behaviors are different: the SST model typically has a well-defined transition location, whereas the SA model does not. Both models are predisposed to delayed activation of turbulence with increasing freestream Mach number. Also, both models can be made to achieve earlier activation of turbulence by increasing their freestream levels, but too high a level can disturb the turbulent solution behavior. The technique of maintaining freestream levels of turbulence without decay in the SST model, introduced elsewhere, is shown here to be useful in reducing grid-dependence of the model's transitional behavior. Both models are demonstrated to be incapable of predicting relaminarization; eddy viscosities remain weakly turbulent in accelerating or laterally-strained boundary layers for which experiment and direct simulations indicate turbulence suppression. The main conclusion is that these models are intended for fully turbulent high Reynolds number computations, and using them for transitional (e.g., low Reynolds number) or relaminarizing flows is not appropriate.
NASA Technical Reports Server (NTRS)
Henderson, W. P.; Huffman, J. K.
1972-01-01
An investigation has been conducted to determine the effects of wing camber and twist on the longitudinal aerodynamic characteristics of a wingbody configuration. Three wings were used each having the same planform (aspect ratio of 2.5 and leading-edge sweep angle of 44 deg.) but differing in amounts of camber and twist (wing design lift coefficient). The wing design lift coefficients were 0, 0.35, and 0.70. The investigation was conducted over a Mach number range from 0.20 to 0.70 at angles of attack up to about 22 deg. The effect of wing strakes on the aerodynamic characteristics of the cambered wings was also studied. A comparison of the experimentally determined aerodynamic characteristics with theoretical estimates is also included.
NASA Technical Reports Server (NTRS)
Gibson, A. F.
1983-01-01
A system of computer programs has been developed to model general three-dimensional surfaces. Surfaces are modeled as sets of parametric bicubic patches. There are also capabilities to transform coordinate to compute mesh/surface intersection normals, and to format input data for a transonic potential flow analysis. A graphical display of surface models and intersection normals is available. There are additional capabilities to regulate point spacing on input curves and to compute surface intersection curves. Internal details of the implementation of this system are explained, and maintenance procedures are specified.
Dynamic Modeling of Starting Aerodynamics and Stage Matching in an Axi-Centrifugal Compressor
NASA Technical Reports Server (NTRS)
Wilkes, Kevin; OBrien, Walter F.; Owen, A. Karl
1996-01-01
A DYNamic Turbine Engine Compressor Code (DYNTECC) has been modified to model speed transients from 0-100% of compressor design speed. The impetus for this enhancement was to investigate stage matching and stalling behavior during a start sequence as compared to rotating stall events above ground idle. The model can simulate speed and throttle excursions simultaneously as well as time varying bleed flow schedules. Results of a start simulation are presented and compared to experimental data obtained from an axi-centrifugal turboshaft engine and companion compressor rig. Stage by stage comparisons reveal the front stages to be operating in or near rotating stall through most of the start sequence. The model matches the starting operating line quite well in the forward stages with deviations appearing in the rearward stages near the start bleed. Overall, the performance of the model is very promising and adds significantly to the dynamic simulation capabilities of DYNTECC.
Comparison of Aerodynamic Resistance Parameterizations and Implications for Dry Deposition Modeling
Nitrogen deposition data used to support the secondary National Ambient Air Quality Standards and critical loads research derives from both measurements and modeling. Data sets with spatial coverage sufficient for regional scale deposition assessments are currently generated fro...
NASA Technical Reports Server (NTRS)
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.
Unsteady Aerodynamic Models for Flight Control of Agile Micro Air Vehicles
2010-08-13
MONITOR’S REPORT NUMBER(S) AFRL-OSR-VA-TR-2011-0251 12. DISTRIBUTION /AVAILABILITY STATEMENT Approved for public release; distribution unlimited 13...the model in [18] must be calibrated to experimental data, and often does not match data it was not specifically calibrated against [9]. The model [41...strategies have a serious limitation in that they need to be calibrated for a particular angle of attack, and are not immediately suitable for
Aerodynamic effects of flexibility in flapping wings.
Zhao, Liang; Huang, Qingfeng; Deng, Xinyan; Sane, Sanjay P
2010-03-06
Recent work on the aerodynamics of flapping flight reveals fundamental differences in the mechanisms of aerodynamic force generation between fixed and flapping wings. When fixed wings translate at high angles of attack, they periodically generate and shed leading and trailing edge vortices as reflected in their fluctuating aerodynamic force traces and associated flow visualization. In contrast, wings flapping at high angles of attack generate stable leading edge vorticity, which persists throughout the duration of the stroke and enhances mean aerodynamic forces. Here, we show that aerodynamic forces can be controlled by altering the trailing edge flexibility of a flapping wing. We used a dynamically scaled mechanical model of flapping flight (Re approximately 2000) to measure the aerodynamic forces on flapping wings of variable flexural stiffness (EI). For low to medium angles of attack, as flexibility of the wing increases, its ability to generate aerodynamic forces decreases monotonically but its lift-to-drag ratios remain approximately constant. The instantaneous force traces reveal no major differences in the underlying modes of force generation for flexible and rigid wings, but the magnitude of force, the angle of net force vector and centre of pressure all vary systematically with wing flexibility. Even a rudimentary framework of wing veins is sufficient to restore the ability of flexible wings to generate forces at near-rigid values. Thus, the magnitude of force generation can be controlled by modulating the trailing edge flexibility and thereby controlling the magnitude of the leading edge vorticity. To characterize this, we have generated a detailed database of aerodynamic forces as a function of several variables including material properties, kinematics, aerodynamic forces and centre of pressure, which can also be used to help validate computational models of aeroelastic flapping wings. These experiments will also be useful for wing design for small
Status of Computational Aerodynamic Modeling Tools for Aircraft Loss-of-Control
NASA Technical Reports Server (NTRS)
Frink, Neal T.; Murphy, Patrick C.; Atkins, Harold L.; Viken, Sally A.; Petrilli, Justin L.; Gopalarathnam, Ashok; Paul, Ryan C.
2016-01-01
A concerted effort has been underway over the past several years to evolve computational capabilities for modeling aircraft loss-of-control under the NASA Aviation Safety Program. A principal goal has been to develop reliable computational tools for predicting and analyzing the non-linear stability & control characteristics of aircraft near stall boundaries affecting safe flight, and for utilizing those predictions for creating augmented flight simulation models that improve pilot training. Pursuing such an ambitious task with limited resources required the forging of close collaborative relationships with a diverse body of computational aerodynamicists and flight simulation experts to leverage their respective research efforts into the creation of NASA tools to meet this goal. Considerable progress has been made and work remains to be done. This paper summarizes the status of the NASA effort to establish computational capabilities for modeling aircraft loss-of-control and offers recommendations for future work.
NASA Astrophysics Data System (ADS)
Mettler, B. F.
2010-09-01
This paper describes a methodology to extract aerial vehicles’ aerodynamic characteristics from visually tracked trajectory data. The technique is being developed to study the aerodynamics of centimeter-scale aircraft and develop flight simulation models. Centimeter-scale aircraft remains a largely unstudied domain of aerodynamics, for which traditional techniques like wind tunnels and computational fluid dynamics have not yet been fully adapted and validated. The methodology takes advantage of recent progress in commercial, vision-based, motion-tracking systems. This system dispenses from on-board navigation sensors and enables indoor flight testing under controlled atmospheric conditions. Given the configuration of retro-reflective markers affixed onto the aerial vehicle, the vehicle’s six degrees-of-freedom motion can be determined in real time. Under disturbance-free conditions, the aerodynamic forces and moments can be determined from the vehicle’s inertial acceleration, and furthermore, for a fixed-wing vehicle, the aerodynamic angles can be plotted from the vehicle’s kinematics. By combining this information, we can determine the temporal evolution of the aerodynamic coefficients, as they change throughout a trajectory. An attractive feature of this technique is that trajectories are not limited to equilibrium conditions but can include non-equilibrium, maneuvering flight. Whereas in traditional wind-tunnel experiments, the operating conditions are set by the experimenter, here, the aerodynamic conditions are driven by the vehicle’s own dynamics. As a result, this methodology could be useful for characterizing the unsteady aerodynamics effects and their coupling with the aircraft flight dynamics, providing insight into aerodynamic phenomena taking place at centimeter scale flight.
A New Aerodynamic Data Dispersion Method for Launch Vehicle Design
NASA Technical Reports Server (NTRS)
Pinier, Jeremy T.
2011-01-01
A novel method for implementing aerodynamic data dispersion analysis is herein introduced. A general mathematical approach combined with physical modeling tailored to the aerodynamic quantity of interest enables the generation of more realistically relevant dispersed data and, in turn, more reasonable flight simulation results. The method simultaneously allows for the aerodynamic quantities and their derivatives to be dispersed given a set of non-arbitrary constraints, which stresses the controls model in more ways than with the traditional bias up or down of the nominal data within the uncertainty bounds. The adoption and implementation of this new method within the NASA Ares I Crew Launch Vehicle Project has resulted in significant increases in predicted roll control authority, and lowered the induced risks for flight test operations. One direct impact on launch vehicles is a reduced size for auxiliary control systems, and the possibility of an increased payload. This technique has the potential of being applied to problems in multiple areas where nominal data together with uncertainties are used to produce simulations using Monte Carlo type random sampling methods. It is recommended that a tailored physics-based dispersion model be delivered with any aerodynamic product that includes nominal data and uncertainties, in order to make flight simulations more realistic and allow for leaner spacecraft designs.
NASA Technical Reports Server (NTRS)
Tweedt, Daniel L.
2014-01-01
Computational Aerodynamic simulations of an 840 ft/sec tip speed, Advanced Ducted Propulsor fan system were performed at five different operating points on the fan operating line, in order to provide detailed internal flow field information for use with fan acoustic prediction methods presently being developed, assessed and validated. The fan system is a sub-scale, lownoise research fan/nacelle model that has undergone extensive experimental testing in the 9- by 15- foot Low Speed Wind Tunnel at the NASA Glenn Research Center, resulting in quality, detailed aerodynamic and acoustic measurement data. Details of the fan geometry, the computational fluid dynamics methods, the computational grids, and various computational parameters relevant to the numerical simulations are discussed. Flow field results for three of the five operating conditions simulated are presented in order to provide a representative look at the computed solutions. Each of the five fan aerodynamic simulations involved the entire fan system, excluding a long core duct section downstream of the core inlet guide vane. As a result, only fan rotational speed and system bypass ratio, set by specifying static pressure downstream of the core inlet guide vane row, were adjusted in order to set the fan operating point, leading to operating points that lie on a fan operating line and making mass flow rate a fully dependent parameter. The resulting mass flow rates are in good agreement with measurement values. The computed blade row flow fields for all five fan operating points are, in general, aerodynamically healthy. Rotor blade and fan exit guide vane flow characteristics are good, including incidence and deviation angles, chordwise static pressure distributions, blade surface boundary layers, secondary flow structures, and blade wakes. Examination of the computed flow fields reveals no excessive boundary layer separations or related secondary-flow problems. A few spanwise comparisons between
The Aerodynamic Drag of Five Models of Side Floats N.A.C.A. Models 51-E, 51-F, 51-G, 51-H, 51-J
NASA Technical Reports Server (NTRS)
House, R O
1938-01-01
The drag of five models of side floats was measured in the N.A.C.A. 7- by 10-foot wind tunnel. The most promising method of reducing the drag of floats indicated by these tests is lowering the angle at which the floats are rigged. The addition of a step to a float does not always increase the drag in the flying range, floats with steps sometimes having lower drag than similar floats without steps. Making the bow chine no higher than necessary might result in a reduction in air drag because of the lower angle of pitch of the chines. Since side floats are used formally to obtain lateral stability when the seaplane is operating on the water at slow speeds or at rest, greater consideration can be given to factors affecting aerodynamic drag than is possible for other types of floats and hulls.
Aerodynamic and propeller performance characteristics of a propfan-powered, semispan model
NASA Technical Reports Server (NTRS)
Levin, Alan D.; Smith, Ronald C.; Wood, Richard D.
1985-01-01
A semispan wing/body model with a powered propeller was tested to provide data on a total powerplant installation drag penalty of advanced propfan-powered aircraft. The test objectives were to determine the total power plant installation drag penalty on a representative propfan aircraft; to study the effect of configuration modifications on the installed powerplant drag; and to determine performance characteristics of an advanced design propeller which was mounted on a representative nacelle in the presence of a wing.
Aerodynamic Characteristics of a Canard and an Outboard-Tail Airplane Model at High Subsonic Speeds
NASA Technical Reports Server (NTRS)
Fournier, Paul G.
1961-01-01
An investigation has been made in the Langley high-speed 7- by 10-foot tunnel through a range of Mach numbers from 0.60 to 0.95 of the static longitudinal and lateral stability and control characteristics of a canard airplane configuration and an outboard-tail configuration. The canard model had a twisted wing with approximately 67 deg of sweepback and an aspect ratio of 2.91 and was tested with three trapezoidal canard surfaces having ratios of exposed area to wing area of 0.032, 0.076, and 0.121. The canard model had a single body-mounted vertical tail. The outboard-tail model had its horizontal- and vertical-tail surfaces mounted on slender bodies attached to the wing tips and located to the rear and outboard of the 67 deg sweptback wing of aspect ratio 1.00. The data, which are presented with limited analysis, provide information at high subsonic speeds on these two types of high-speed airplanes which have previously been tested at supersonic speeds and reported in NACA RM L58BO7 and NACA RM L58E20.
NASA Technical Reports Server (NTRS)
West, F. E., Jr.
1959-01-01
The longitudinal aerodynamic characteristics of a wing-body-horizontal-tail configuration designed for efficient performance at transonic speeds has been investigated at Mach numbers from 0.80 to 1.03 in the Langley 16-foot transonic tunnel. The effect of adding an outboard leading-edge chord-extension to the highly tapered 45 deg. swept wing was also obtained. The average Reynolds number for this investigation was 6.7 x 10(exp 6) based on the wing mean aerodynamic chord. The relatively low tail placement as well as the addition of a chord-extension achieved some alleviation of the pitchup tendencies of the wing-fuselage configuration. The maximum trimmed lift-drag ratio was 16.5 up to a Mach number of 0.9, with the moment center located at the quarter-chord point of the mean aerodynamic chord. For the untrimmed case, the maximum lift-drag ratio was approximately 19.5 up to a Mach number of 0.9.
2015-01-05
canopies , often referred to as parafoils. No numerical studies, however, have fully investigated the 3-D aerodynamic performance of these bleed-air actuators...Simulation results are presented for a finite span, ram-air canopy geometry and several configurations amenable for comparison with wind tunnel
On the Aerodynamic Characteristics over Idealized Two-Dimensional Urban Street Canyon Models
NASA Astrophysics Data System (ADS)
Leung, K. K.; Liu, C. H.
2012-04-01
There are numerous anthropogenic pollutant sources in the atmospheric boundary layer (ABL) nowadays, which mainly attributed to human activities in urban areas. Hence, how urban morphology affects the heat and mass transfer in built environment is a popular research problem in the urban climate community. However, our understanding of street-level transport processes is rather limited. Laboratory experiments often serve as complementary solutions to modeling results. Although there are laboratory results available for the mass transfer over idealized urban roughness, the transport processes are not examined in details. In this paper, we attempt to demystify the pollutant removal mechanism from urban areas to the urban ABL. Laboratory measurements, which were conducted in the wind tunnel in Mechanical Engineering, The University of Hong Kong, and computational fluid dynamics (CFD) is used concurrently. The spatial air pollutant transport from the street region to the urban ABL was represented by means of water evaporation method from the soaked filter paper applied on the surfaces of the building facades and ground surface. Street canyon models of building-height-to-street-width (aspect) ratios in the range of 0.125 to 2 are carried out. The local mass transfer velocity along the street canyons was measured and archived a good comparison with the outside literature. Besides, both the laboratory and CFD results show that the pollutant removal from 2D street canyons increases with decreasing ARs. It arrives a local maximum then decreases thereafter. In the comparison between laboratory and CFD results, the difference in the size of the street canyon models, also known as scaling effects, is needed to be considered. Therefore, despite of representing the transfer behavior by the local pollutant exchange rate, scaled local/overall pollutant removal coefficient is proposed for a comparison of pollutant removal performance in a more reasonable manner. Such effect is found
Aerodynamic characteristics of a large scale model with a swept wing and augmented jet flap
NASA Technical Reports Server (NTRS)
Falarski, M. D.; Koenig, D. G.
1971-01-01
Data of tests of a large-scale swept augmentor wing model in the 40- by 80-foot wind tunnel are presented. The data includes longitudinal characteristics with and without a horizontal tail as well as results of preliminary investigation of lateral-directional characteristics. The augmentor flap deflection was varied from 0 deg to 70.6 deg at isentropic jet thrust coefficients of 0 to 1.47. The tests were made at a Reynolds number from 2.43 to 4.1 times one million.
Aerodynamic Analyses Requiring Advanced Computers, Part 1
NASA Technical Reports Server (NTRS)
1975-01-01
Papers are presented which deal with results of theoretical research on aerodynamic flow problems requiring the use of advanced computers. Topics discussed include: viscous flows, boundary layer equations, turbulence modeling and Navier-Stokes equations, and internal flows.
NASA Technical Reports Server (NTRS)
Banks, Daniel W.; Kelley, Henry L.
2000-01-01
Two large-scale, two-dimensional helicopter tail boom models were used to determine the effects of passive venting on boom down loads and side forces in hovering crosswind conditions. The models were oval shaped and trapezoidal shaped. Completely porous and solid configurations, partial venting in various symmetric and asymmetric configurations, and strakes were tested. Calculations were made to evaluate the trends of venting and strakes on power required when applied to a UH-60 class helicopter. Compared with the UH-60 baseline, passive venting reduced side force but increased down load at flow conditions representing right sideward flight. Selective asymmetric venting resulted in side force benefits close to the fully porous case. Calculated trends on the effects of venting on power required indicated that the high asymmetric oval configuration was the most effective venting configuration for side force reduction, and the high asymmetric with a single strake was the most effective for overall power reduction. Also, curves of side force versus flow angle were noticeable smoother for the vented configurations compared with the solid baseline configuration; this indicated a potential for smoother flight in low-speed crosswind conditions.
1979-02-01
desired properties. To this end, we shall make use of ex- perimental evidence, as obtained in wind or water tunnels by various flow visualization...based on the results of number of tests carried out in a water and various wind tunnels, as reported by H. Werl6 in Ref. 3. Vortex breakdown is found...of a triangular wing Rogachev, G.V. moving close to the earth ’ s surface. AD 785154, FTD-HC-23-1802-74. 25 Fox, C.H. Prediction of lift and drag for
Flow aerodynamics modeling of an MHD swirl combustor - Calculations and experimental verification
NASA Technical Reports Server (NTRS)
Gupta, A. K.; Beer, J. M.; Louis, J. F.; Busnaina, A. A.; Lilley, D. G.
1981-01-01
The paper describes a computer code for calculating the flow dynamics of a constant-density flow in the second-stage trumpet shaped nozzle section of a two-stage MHD swirl combustor for application to a disk generator. The primitive pressure-velocity variable, finite-difference computer code has been developed for the computation of inert nonreacting turbulent swirling flows in an axisymmetric MHD model swirl combustor. The method and program involve a staggered grid system for axial and radial velocities, and a line relaxation technique for the efficient solution of the equations. The code produces as output the flow field map of the nondimensional stream function, axial and swirl velocity. It was found that the best location for seed injection to obtain a uniform distribution at the combustor exit is in the central location for seed injected at the entrance to the second stage combustor.
NASA Technical Reports Server (NTRS)
Choo, Yung K.; Slater, John W.; Vickerman, Mary B.; VanZante, Judith F.; Wadel, Mary F. (Technical Monitor)
2002-01-01
Issues associated with analysis of 'icing effects' on airfoil and wing performances are discussed, along with accomplishments and efforts to overcome difficulties with ice. Because of infinite variations of ice shapes and their high degree of complexity, computational 'icing effects' studies using available software tools must address many difficulties in geometry acquisition and modeling, grid generation, and flow simulation. The value of each technology component needs to be weighed from the perspective of the entire analysis process, from geometry to flow simulation. Even though CFD codes are yet to be validated for flows over iced airfoils and wings, numerical simulation, when considered together with wind tunnel tests, can provide valuable insights into 'icing effects' and advance our understanding of the relationship between ice characteristics and their effects on performance degradation.
NASA Technical Reports Server (NTRS)
Fournier, P. G.; Goodson, K. W.
1974-01-01
A low-speed investigation was conducted over an angle-of-attack range from about -4 deg to 20 deg in the Langley V/STOL tunnel to determine the effects of a double-slotted flap, high-lift system on the aerodynamic characteristics of a 42 deg swept high-wing model having a supercritical airfoil. The wing had an aspect ratio of 6.78 and a taper ratio of 0.36; the double-slotted flap consisted of a 35-percent-chord flap with a 15-percent-chord vane. The model was tested with a 15-percent-chord leading-edge slat.
NASA Technical Reports Server (NTRS)
Campbell, R. L.
1982-01-01
Tests were conducted in the Langley High-Speed 7- by 10-Foot Tunnel using a 1/10-scale model of an executive jet to examine the effects of the nacelles on the wing pressures and model longitudinal aerodynamic characteristics. For the present investigation, each wing panel was modified with a simulated, partial-chord, laminar-flow-control glove. Horizontal-tail effects were also briefly examined. The tests covered a range of Mach numbers from 0.40 to 0.82 and lift coefficients from 0.20 to 0.55. Oil-flow photographs of the wing at selected conditions are included.
NASA Astrophysics Data System (ADS)
ELGAMMI, MOUTAZ; SANT, TONIO
2016-09-01
This paper investigates a new approach to model the stochastic variations in the aerodynamic loads on yawed wind turbines experienced at high angles of attack. The method applies the one-dimensional Langevin equation in conjunction with known mean and standard deviation values for the lift and drag data. The method is validated using the experimental data from the NREL Phase VI rotor in which the mean and standard deviation values for the lift and drag are derived through the combined use of blade pressure measurements and a free-wake vortex model. Given that direct blade pressure measurements are used, 3D flow effects arising from the co-existence of dynamic stall and stall delay are taken into account. The model is an important step towards verification of several assumptions characterized as the estimated standard deviation, Gaussian white noise of the data and the estimated drift and diffusion coefficients of the Langevin equation. The results using the proposed assumptions lead to a good agreement with measurements over a wide range of operating conditions. This provides motivation to implement a general fully independent theoretical stochastic model within a rotor aerodynamics model, such as the free-wake vortex or blade-element momentum code, whereby the mean lift and drag coefficients can be estimated using 2D aerofoil data with correction models for 3D dynamic stall and stall delay phenomena, while the corresponding standard derivations are estimated through CFD.
Active Control of Aerodynamic Noise Sources
NASA Technical Reports Server (NTRS)
Reynolds, Gregory A.
2001-01-01
Aerodynamic noise sources become important when propulsion noise is relatively low, as during aircraft landing. Under these conditions, aerodynamic noise from high-lift systems can be significant. The research program and accomplishments described here are directed toward reduction of this aerodynamic noise. Progress toward this objective include correction of flow quality in the Low Turbulence Water Channel flow facility, development of a test model and traversing mechanism, and improvement of the data acquisition and flow visualization capabilities in the Aero. & Fluid Dynamics Laboratory. These developments are described in this report.
Numerical modeling of aerodynamics of airfoils of micro air vehicles in gusty environment
NASA Astrophysics Data System (ADS)
Gopalan, Harish
The superior flight characteristics exhibited by birds and insects can be taken as a prototype of the most perfect form of flying machine ever created. The design of Micro Air Vehicles (MAV) which tries mimic the flight of birds and insects has generated a great deal of interest as the MAVs can be utilized for a number of commercial and military operations which is usually not easily accessible by manned motion. The size and speed of operation of a MAV results in low Reynolds number flight, way below the flying conditions of a conventional aircraft. The insensitivity to wind shear and gust is one of the required factors to be considered in the design of airfoil for MAVs. The stability of flight under wind shear is successfully accomplished in the flight of birds and insects, through the flapping motion of their wings. Numerous studies which attempt to model the flapping motion of the birds and insects have neglected the effect of wind gust on the stability of the motion. Also sudden change in flight conditions makes it important to have the ability to have an instantaneous change of the lift force without disturbing the stability of the MAV. In the current study, two dimensional rigid airfoil, undergoing flapping motion is studied numerically using a compressible Navier-Stokes solver discretized using high-order finite difference schemes. The high-order schemes in space and in time are needed to keep the numerical solution economic in terms of computer resources and to prevent vortices from smearing. The numerical grid required for the computations are generated using an inverse panel method for the streamfunction and potential function. This grid generating algorithm allows the creation of single-block orthogonal H-grids with ease of clustering anywhere in the domain and the easy resolution of boundary layers. The developed numerical algorithm has been validated successfully against benchmark problems in computational aeroacoustics (CAA), and unsteady viscous
Transpiration Control Of Aerodynamics Via Porous Surfaces
NASA Technical Reports Server (NTRS)
Banks, Daniel W.; Wood, Richard M.; Bauer, Steven X. S.
1993-01-01
Quasi-active porous surface used to control pressure loading on aerodynamic surface of aircraft or other vehicle, according to proposal. In transpiration control, one makes small additions of pressure and/or mass to cavity beneath surface of porous skin on aerodynamic surface, thereby affecting rate of transpiration through porous surface. Porous skin located on forebody or any other suitable aerodynamic surface, with cavity just below surface. Device based on concept extremely lightweight, mechanically simple, occupies little volume in vehicle, and extremely adaptable.
Aerodynamic Simulation of the MEXICO Rotor
NASA Astrophysics Data System (ADS)
Herraez, I.; Medjroubi, W.; Stoevesandt, B.; Peinke, J.
2014-12-01
CFD (Computational Fluid Dynamics) simulations are a very promising method for predicting the aerodynamic behavior of wind turbines in an inexpensive and accurate way. One of the major drawbacks of this method is the lack of validated models. As a consequence, the reliability of numerical results is often difficult to assess. The MEXICO project aimed at solving this problem by providing the project partners with high quality measurements of a 4.5 meters rotor diameter wind turbine operating under controlled conditions. The large measurement data-set allows the validation of all kind of aerodynamic models. This work summarizes our efforts for validating a CFD model based on the open source software OpenFoam. Both steady- state and time-accurate simulations have been performed with the Spalart-Allmaras turbulence model for several operating conditions. In this paper we will concentrate on axisymmetric inflow for 3 different wind speeds. The numerical results are compared with pressure distributions from several blade sections and PIV-flow data from the near wake region. In general, a reasonable agreement between measurements the and our simulations exists. Some discrepancies, which require further research, are also discussed.
Uncertainty in Computational Aerodynamics
NASA Technical Reports Server (NTRS)
Luckring, J. M.; Hemsch, M. J.; Morrison, J. H.
2003-01-01
An approach is presented to treat computational aerodynamics as a process, subject to the fundamental quality assurance principles of process control and process improvement. We consider several aspects affecting uncertainty for the computational aerodynamic process and present a set of stages to determine the level of management required to meet risk assumptions desired by the customer of the predictions.
NASA Technical Reports Server (NTRS)
Smith, Mark S.; Bui, Trong T.; Garcia, Christian A.; Cumming, Stephen B.
2016-01-01
A pair of compliant trailing edge flaps was flown on a modified GIII airplane. Prior to flight test, multiple analysis tools of various levels of complexity were used to predict the aerodynamic effects of the flaps. Vortex lattice, full potential flow, and full Navier-Stokes aerodynamic analysis software programs were used for prediction, in addition to another program that used empirical data. After the flight-test series, lift and pitching moment coefficient increments due to the flaps were estimated from flight data and compared to the results of the predictive tools. The predicted lift increments matched flight data well for all predictive tools for small flap deflections. All tools over-predicted lift increments for large flap deflections. The potential flow and Navier-Stokes programs predicted pitching moment coefficient increments better than the other tools.
NASA Technical Reports Server (NTRS)
Enomoto, F.; Keller, P.
1984-01-01
The Computer Aided Design (CAD) system's common geometry database was used to generate input for theoretical programs and numerically controlled (NC) tool paths for wind tunnel part fabrication. This eliminates the duplication of work in generating separate geometry databases for each type of analysis. Another advantage is that it reduces the uncertainty due to geometric differences when comparing theoretical aerodynamic data with wind tunnel data. The system was adapted to aerodynamic research by developing programs written in Design Analysis Language (DAL). These programs reduced the amount of time required to construct complex geometries and to generate input for theoretical programs. Certain shortcomings of the Design, Drafting, and Manufacturing (DDM) software limited the effectiveness of these programs and some of the Calma NC software. The complexity of aircraft configurations suggests that more types of surface and curve geometry should be added to the system. Some of these shortcomings may be eliminated as improved versions of DDM are made available.
NASA Astrophysics Data System (ADS)
Bragg, M. B.; Broeren, A. P.; Blumenthal, L. A.
2005-07-01
Past research on airfoil aerodynamics in icing are reviewed. This review emphasizes the time period after the 1978 NASA Lewis workshop that initiated the modern icing research program at NASA and the current period after the 1994 ATR accident where aerodynamics research has been more aircraft safety focused. Research pre-1978 is also briefly reviewed. Following this review, our current knowledge of iced airfoil aerodynamics is presented from a flowfield-physics perspective. This article identifies four classes of ice accretions: roughness, horn ice, streamwise ice, and spanwise-ridge ice. For each class, the key flowfield features such as flowfield separation and reattachment are discussed and how these contribute to the known aerodynamic effects of these ice shapes. Finally Reynolds number and Mach number effects on iced-airfoil aerodynamics are summarized.
An Overview of Ares-I CFD Ascent Aerodynamic Data Development And Analysis Based on USM3D
NASA Technical Reports Server (NTRS)
Abdol-Hamid, Khaled S.; Ghaffari, Farhad; Parlette, Edward B.
2011-01-01
An overview of the computational results obtained from the NASA Langley developed unstructured grid, Reynolds-averaged Navier-Stokes flow solver USM3D, in support of the Ares-I project within the NASA s Constellation program, are presented. The numerical data are obtained for representative flow conditions pertinent to the ascent phase of the trajectory at both wind tunnel and flight Reynolds number without including any propulsion effects. The USM3D flow solver has been designated to have the primary role within the Ares-I project in developing the computational aerodynamic data for the vehicle while other flow solvers, namely OVERFLOW and FUN3D, have supporting roles to provide complementary results for fewer cases as part of the verification process to ensure code-to-code solution consistency. Similarly, as part of the solution validation efforts, the predicted numerical results are correlated with the aerodynamic wind tunnel data that have been generated within the project in the past few years. Sample aerodynamic results and the processes established for the computational solution/data development for the evolving Ares-I design cycles are presented.
NASA Technical Reports Server (NTRS)
Kuhn, Richard E; Draper, John W
1956-01-01
This report presents the results of an investigation conducted in the Langley 300 mph 7- by 10-foot wind tunnel for the purpose of determining the aerodynamic characteristics of a model wing-propeller combination, and of the wing and propeller separately at angles of attack up to 90 degrees. The tests covered thrust coefficients corresponding to free-stream velocities from zero forward speed to the normal range of cruising speeds. The results indicate that increasing the thrust coefficient increases the angle of attack for maximum lift and greatly diminishes the usual reduction in lift above the angle of attack for maximum lift.
NASA Technical Reports Server (NTRS)
Nelson, D. P.
1981-01-01
A graphical presentation of the aerodynamic data acquired during coannular nozzle performance wind tunnel tests is given. The graphical data consist of plots of nozzle gross thrust coefficient, fan nozzle discharge coefficient, and primary nozzle discharge coefficient. Normalized model component static pressure distributions are presented as a function of primary total pressure, fan total pressure, and ambient static pressure for selected operating conditions. In addition, the supersonic cruise configuration data include plots of nozzle efficiency and secondary-to-fan total pressure pumping characteristics. Supersonic and subsonic cruise data are given.
Workshop on Aircraft Surface Representation for Aerodynamic Computation
NASA Technical Reports Server (NTRS)
Gregory, T. J. (Editor); Ashbaugh, J. (Editor)
1980-01-01
Papers and discussions on surface representation and its integration with aerodynamics, computers, graphics, wind tunnel model fabrication, and flow field grid generation are presented. Surface definition is emphasized.
NASA Technical Reports Server (NTRS)
Dring, R. P.; Joslyn, H. D.; Blair, M. F.
1987-01-01
A combined experimental and analytical program was conducted to examine the effects of inlet turbulence and airfoil heat transfer. The experimental portion of the study was conducted in a large-scale (approx. 5X engine), ambient temperature, rotating turbine model configured in both single-stage and stage-and-a-half arrangements. Heat transfer measurements were obtained using low-conductivity airfoils with miniature thermocouples welded to a thin, electrically heated surface skin. Heat transfer data were acquired for various combinations of low or high inlet turbulence intensity, flow coefficient, first stator-rotor axial spacing, Reynolds number and relative circumferential position of the first and second stators. Aerodynamic measurements obtained include distributions of the mean and fluctuating velocities at the turbine inlet and, for each airfoil row, midspan airfoil surface pressures and circumferential distributions of the downstream steady state pressures and fluctuating velocities. Results include airfoil heat transfer predictions produced using existing 2-D boundary layer computation schemes and an examination of solutions of the unsteady boundary layer equations.
NASA Technical Reports Server (NTRS)
Commo, Sean A. (Inventor); Lynn, Keith C. (Inventor); Landman, Drew (Inventor); Acheson, Michael J. (Inventor)
2016-01-01
An In-Situ Load System for calibrating and validating aerodynamic properties of scaled aircraft in ground-based aerospace testing applications includes an assembly having upper and lower components that are pivotably interconnected. A test weight can be connected to the lower component to apply a known force to a force balance. The orientation of the force balance can be varied, and the measured forces from the force balance can be compared to applied loads at various orientations to thereby develop calibration factors.
NASA Technical Reports Server (NTRS)
Tweedt, Daniel L.
2014-01-01
Computational Aerodynamic simulations of a 1215 ft/sec tip speed transonic fan system were performed at five different operating points on the fan operating line, in order to provide detailed internal flow field information for use with fan acoustic prediction methods presently being developed, assessed and validated. The fan system is a sub-scale, low-noise research fan/nacelle model that has undergone extensive experimental testing in the 9- by 15-foot Low Speed Wind Tunnel at the NASA Glenn Research Center. Details of the fan geometry, the computational fluid dynamics methods, the computational grids, and various computational parameters relevant to the numerical simulations are discussed. Flow field results for three of the five operating points simulated are presented in order to provide a representative look at the computed solutions. Each of the five fan aerodynamic simulations involved the entire fan system, which for this model did not include a split flow path with core and bypass ducts. As a result, it was only necessary to adjust fan rotational speed in order to set the fan operating point, leading to operating points that lie on a fan operating line and making mass flow rate a fully dependent parameter. The resulting mass flow rates are in good agreement with measurement values. Computed blade row flow fields at all fan operating points are, in general, aerodynamically healthy. Rotor blade and fan exit guide vane flow characteristics are good, including incidence and deviation angles, chordwise static pressure distributions, blade surface boundary layers, secondary flow structures, and blade wakes. Examination of the flow fields at all operating conditions reveals no excessive boundary layer separations or related secondary-flow problems.
NASA Technical Reports Server (NTRS)
Fischetti, Thomas L.
1958-01-01
An investigation has been made in the Langley 8-foot transonic tunnels on the aerodynamic characteristics of a 0.15-scale model of the North American Aviation 255-inch fin-stabilized external store over a maximum Mach number range of 0.60 to 1.2 and on the effects of mounting lugs, of fin orientation, of fin aspect ratio, and of fixed-transition. The Reynolds number (based on a body length of 37.50 inches) varied from 9.8 x 10(exp 6) to 13.1 x 10(exp 6). The results indicate that the static margin of the finned store at low lift coefficients was only 9 percent of body length at subsonic Mach numbers and was reduced to zero at a Mach number of 1.0, Increasing the fin aspect ratio from 1.82 to 2.41 increased the subsonic static margin to 18 percent and provided a minimum margin of 9 percent near a Mach number of l.O. Store mounting lugs or fin orientation had only small effects on the aerodynamic characteristics of the basic store.
A tomographic technique for aerodynamics at transonic speeds
NASA Technical Reports Server (NTRS)
Lee, G.
1985-01-01
Computer aided tomography (CAT) provides a means of noninvasively measuring the air density distribution around an aerodynamic model. This technique is global in that a large portion of the flow field can be measured. A test of the applicability of CAT to transonic velocities was studied. A hemispherical-nose cylinder afterbody model was tested at a Mach number of 0.8 with a new laser holographic interferometer at the 2- by 2-Foot Transonic Wind Tunnel. Holograms of the flow field were taken and were reconstructed into interferograms. The fringe distribution (a measure of the local densities) was digitized for subsequent data reduction. A computer program based on the Fourier-transform technique was developed to convert the fringe distribution into three-dimensional densities around the model. Theoretical aerodynamic densities were calculated for evaluating and assessing the accuracy of the data obtained from the tomographic method.
NASA Technical Reports Server (NTRS)
Thornton, D. E.
1976-01-01
Tests were conducted in a 14 foot transonic wind tunnel to examine the feasibility of the auxiliary aerodynamic data system (AADS) for determining angles of attack and sideslip during boost flight. The model used was a 0.07 scale replica of the external tank forebody consisting of the nose portion and a 60 inch (full scale) cylindrical section of the ogive cylinder tangency point. The model terminated in a blunt base with a 320.0 inch diameter at external tank (ET) station 1120.37. Pressure data were obtained from five pressure orifices (one total and four statics) on the nose probe, and sixteen surface static pressure orifices along the ET forebody.
2012-09-01
14 Figure 9. UCAV 1303 Model with Dimensions (Inches). From [1]. ..............................15 Figure 10. A Small Rubber Tube Placed over the...DC motor. A third motor is enclosed in a waterproofed mechanism which supports roll motions. The model is supported in the inverted position in order...entire system, from tunnel velocity to model motion, is driven by a PC based LabVIEW software. B. THE UCAV 1303 MODEL In order to perform a flow
Introduction. Computational aerodynamics.
Tucker, Paul G
2007-10-15
The wide range of uses of computational fluid dynamics (CFD) for aircraft design is discussed along with its role in dealing with the environmental impact of flight. Enabling technologies, such as grid generation and turbulence models, are also considered along with flow/turbulence control. The large eddy simulation, Reynolds-averaged Navier-Stokes and hybrid turbulence modelling approaches are contrasted. The CFD prediction of numerous jet configurations occurring in aerospace are discussed along with aeroelasticity for aeroengine and external aerodynamics, design optimization, unsteady flow modelling and aeroengine internal and external flows. It is concluded that there is a lack of detailed measurements (for both canonical and complex geometry flows) to provide validation and even, in some cases, basic understanding of flow physics. Not surprisingly, turbulence modelling is still the weak link along with, as ever, a pressing need for improved (in terms of robustness, speed and accuracy) solver technology, grid generation and geometry handling. Hence, CFD, as a truly predictive and creative design tool, seems a long way off. Meanwhile, extreme practitioner expertise is still required and the triad of computation, measurement and analytic solution must be judiciously used.
Aerodynamics of a rolling airframe missile
NASA Astrophysics Data System (ADS)
Tisserand, L. E.
1981-05-01
For guidance-related reasons, there is considerable interest in rolling missiles having single-plane steering capability. To aid the aerodynamic design of these airframes, a unique investigation into the aerodynamics of a rolling, steering missile has been carried out. It represents the first known attempt to measure in a wind tunnel the aerodynamic forces and moments that act on a spinning body-canard-tail configuration that exercises canard steering in phase with body roll position. Measurements were made with the model spinning at steady-state roll rates ranging from 15 to 40 Hz over an angle-of-attack range up to about 16 deg. This short, exploratory investigation has demonstrated that a better understanding and a more complete definition of the aerodynamics of rolling, steering vehicles can be developed by way of simulative wind-tunnel testing.
Uniaxial aerodynamic attitude control of artificial satellites
NASA Technical Reports Server (NTRS)
Sazonov, V. V.
1983-01-01
Within the context of a simple mechanical model the paper examines the movement of a satellite with respect to the center of masses under conditions of uniaxial aerodynamic attitude control. The equations of motion of the satellite take account of the gravitational and restorative aerodynamic moments. It is presumed that the aerodynamic moment is much larger than the gravitational, and the motion equations contain a large parameter. A two-parameter integrated surface of these equations is constructed in the form of formal series in terms of negative powers of the large parameter, describing the oscillations and rotations of the satellite about its lengthwise axis, approximately oriented along the orbital tangent. It is proposed to treat such movements as nominal undisturbed motions of the satellite under conditions of aerodynamic attitude control. A numerical investigation is made for the above integrated surface.
NASA Technical Reports Server (NTRS)
Grunwald, Kalman J.
1961-01-01
Results are presented of a wind-tunnel investigation of the aerodynamic stability, control, and performance characteristics of a model of a four-propeller tilt-wing VTOL airplane employing flaps and speed brakes through the transition speed range. The results indicate that the wing was stalled for steady level flight for all conditions of the investigation; however, the flapped configuration did produce a higher maximum lift. The effectiveness of the flap in delaying the stall in the present investigation was not as great as in some previous investigations because the flap used was smaller than that used previously. The wing stall resulted in an appreciable reduction of aileron effectiveness during the transition. Out of ground effect the low horizontal tail did not appear to be in an adverse flow field as had been expected and showed no erratic changes in effectiveness; however, in ground effect a large nose-down moment was experienced by the model. In general, the lateral aerodynamic data indicate that the configuration is directionally stable and possesses positive dihedral effect throughout the transition, and the data show no signs of erratic flow at the vertical tails.
Steady incompressible variable thickness shear layer aerodynamics
NASA Technical Reports Server (NTRS)
Chi, M. R.
1976-01-01
A shear flow aerodynamic theory for steady incompressible flows is presented for both the lifting and non lifting problems. The slow variation of the boundary layer thickness is considered. The slowly varying behavior is treated by using multitime scales. The analysis begins with the elementary wavy wall problem and, through Fourier superpositions over the wave number space, the shear flow equivalents to the aerodynamic transfer functions of classical potential flow are obtained. The aerodynamic transfer functions provide integral equations which relate the wall pressure and the upwash. Computational results are presented for the pressure distribution, the lift coefficient, and the center of pressure travel along a two dimensional flat plate in a shear flow. The aerodynamic load is decreased by the shear layer, compared to the potential flow. The variable thickness shear layer decreases it less than the uniform thickness shear layer based upon equal maximum shear layer thicknesses.
Aerodynamic Parameter Estimation for the X-43A (Hyper-X) from Flight Data
NASA Technical Reports Server (NTRS)
Morelli, Eugene A.; Derry, Stephen D.; Smith, Mark S.
2005-01-01
Aerodynamic parameters were estimated based on flight data from the third flight of the X-43A hypersonic research vehicle, also called Hyper-X. Maneuvers were flown using multiple orthogonal phase-optimized sweep inputs applied as simultaneous control surface perturbations at Mach 8, 7, 6, 5, 4, and 3 during the vehicle descent. Aerodynamic parameters, consisting of non-dimensional longitudinal and lateral stability and control derivatives, were estimated from flight data at each Mach number. Multi-step inputs at nearly the same flight conditions were also flown to assess the prediction capability of the identified models. Prediction errors were found to be comparable in magnitude to the modeling errors, which indicates accurate modeling. Aerodynamic parameter estimates were plotted as a function of Mach number, and compared with estimates from the pre-flight aerodynamic database, which was based on wind-tunnel tests and computational fluid dynamics. Agreement between flight estimates and values computed from the aerodynamic database was excellent overall.
Aerodynamic Design Using Neural Networks
NASA Technical Reports Server (NTRS)
Rai, Man Mohan; Madavan, Nateri K.
2003-01-01
The design of aerodynamic components of aircraft, such as wings or engines, involves a process of obtaining the most optimal component shape that can deliver the desired level of component performance, subject to various constraints, e.g., total weight or cost, that the component must satisfy. Aerodynamic design can thus be formulated as an optimization problem that involves the minimization of an objective function subject to constraints. A new aerodynamic design optimization procedure based on neural networks and response surface methodology (RSM) incorporates the advantages of both traditional RSM and neural networks. The procedure uses a strategy, denoted parameter-based partitioning of the design space, to construct a sequence of response surfaces based on both neural networks and polynomial fits to traverse the design space in search of the optimal solution. Some desirable characteristics of the new design optimization procedure include the ability to handle a variety of design objectives, easily impose constraints, and incorporate design guidelines and rules of thumb. It provides an infrastructure for variable fidelity analysis and reduces the cost of computation by using less-expensive, lower fidelity simulations in the early stages of the design evolution. The initial or starting design can be far from optimal. The procedure is easy and economical to use in large-dimensional design space and can be used to perform design tradeoff studies rapidly. Designs involving multiple disciplines can also be optimized. Some practical applications of the design procedure that have demonstrated some of its capabilities include the inverse design of an optimal turbine airfoil starting from a generic shape and the redesign of transonic turbines to improve their unsteady aerodynamic characteristics.
The aerodynamics of insect flight.
Sane, Sanjay P
2003-12-01
The flight of insects has fascinated physicists and biologists for more than a century. Yet, until recently, researchers were unable to rigorously quantify the complex wing motions of flapping insects or measure the forces and flows around their wings. However, recent developments in high-speed videography and tools for computational and mechanical modeling have allowed researchers to make rapid progress in advancing our understanding of insect flight. These mechanical and computational fluid dynamic models, combined with modern flow visualization techniques, have revealed that the fluid dynamic phenomena underlying flapping flight are different from those of non-flapping, 2-D wings on which most previous models were based. In particular, even at high angles of attack, a prominent leading edge vortex remains stably attached on the insect wing and does not shed into an unsteady wake, as would be expected from non-flapping 2-D wings. Its presence greatly enhances the forces generated by the wing, thus enabling insects to hover or maneuver. In addition, flight forces are further enhanced by other mechanisms acting during changes in angle of attack, especially at stroke reversal, the mutual interaction of the two wings at dorsal stroke reversal or wing-wake interactions following stroke reversal. This progress has enabled the development of simple analytical and empirical models that allow us to calculate the instantaneous forces on flapping insect wings more accurately than was previously possible. It also promises to foster new and exciting multi-disciplinary collaborations between physicists who seek to explain the phenomenology, biologists who seek to understand its relevance to insect physiology and evolution, and engineers who are inspired to build micro-robotic insects using these principles. This review covers the basic physical principles underlying flapping flight in insects, results of recent experiments concerning the aerodynamics of insect flight, as well
ERIC Educational Resources Information Center
Weltner, Klaus
1990-01-01
Describes some experiments showing both qualitatively and quantitatively that aerodynamic lift is a reaction force. Demonstrates reaction forces caused by the acceleration of an airstream and the deflection of an airstream. Provides pictures of demonstration apparatus and mathematical expressions. (YP)
NASA Technical Reports Server (NTRS)
Fournier, P. G.; Sleeman, W. C., Jr.
1972-01-01
A low speed wind tunnel test was conducted in the Langley V/STOL tunnel to determine the static longitudinal and lateral stability characteristics of a general research model which simulated an advance configuration for a commercial transport airplane with a T tail. The model had a 42 deg swept, aspect ratio 6.78 wing with a supercritical airfoil and a high lift system which consisted of a leading edge slat and a double slotted flap. Various slat and flap deflection combinations represented clean, take off, and landing configurations. Effects on the longitudinal and lateral aerodynamic characteristics were determined for two flow through, simulated engine nacelles located on the sides of the fuselage near the rear of the model.
NASA Technical Reports Server (NTRS)
Horstman, Raymond H.
1992-01-01
Aerodynamic flow achieved by adding fixed fairings to butterfly valve. When valve fully open, fairings align with butterfly and reduce wake. Butterfly free to turn, so valve can be closed, while fairings remain fixed. Design reduces turbulence in flow of air in internal suction system. Valve aids in development of improved porous-surface boundary-layer control system to reduce aerodynamic drag. Applications primarily aerospace. System adapted to boundary-layer control on high-speed land vehicles.
Numerical investigation of wind turbine and wind farm aerodynamics
NASA Astrophysics Data System (ADS)
Selvaraj, Suganthi
A numerical method based on the solution of Reynolds Averaged Navier Stokes equations and actuator disk representation of turbine rotor is developed and implemented in the OpenFOAM software suite for aerodynamic analysis of horizontal axis wind turbines (HAWT). The method and the implementation are validated against the 1-D momentum theory, the blade element momentum theory and against experimental data. The model is used for analyzing aerodynamics of a novel dual rotor wind turbine concept and wind farms. Horizontal axis wind turbines suffer from aerodynamic inefficiencies in the blade root region (near the hub) due to several non-aerodynamic constraints (e.g., manufacturing, transportation, cost, etc.). A new dual-rotor wind turbine (DRWT) concept is proposed that aims at mitigating these losses. A DRWT is designed using an existing turbine rotor for the main rotor (Risoe turbine and NREL 5 MW turbine), while the secondary rotor is designed using a high lift to drag ratio airfoil (the DU 96 airfoil from TU Delft). The numerical aerodynamic analysis method developed as a part of this thesis is used to optimize the design. The new DRWT design gives an improvement of about 7% in aerodynamic efficiency over the single rotor turbine. Wind turbines are typically deployed in clusters called wind farms. HAWTs also suffer from aerodynamic losses in a wind farm due to interactions with wind turbine wakes. An interesting mesoscale meteorological phenomenon called "surface flow convergence" believed to be caused by wind turbine arrays is investigated using the numerical method developed here. This phenomenon is believed to be caused by the pressure gradient set up by wind turbines operating in close proximity in a farm. A conceptual/hypothetical wind farm simulation validates the hypothesis that a pressure gradient is setup in wind farms due to turbines and that it can cause flow veering of the order of 10 degrees. Simulations of a real wind farm (Story County) are also
NASA Technical Reports Server (NTRS)
Parkinson, John B; Olson, Roland E; House, Rufus O
1939-01-01
Three models of V-bottom floats for twin-float seaplanes (N.A.C.A. models 57-A, 57-B, and 57-C) having angles of dead rise of 20 degrees, 25 degrees, and thirty degrees, respectively, were tested in the N.A.C.A. tank and in the N.A.C.A. 7- by 10-foot wind tunnel. Within the range investigated, the effect of angle of dead rise on water resistance was found to be negligible at speeds up to and including the hump speed, and water resistance was found to increase with angle of dead rise at planing speeds. The height of the spray at the hump speed decreased with increase in angle of dead rise and the aerodynamic drag increased with dead rise. Lengthening the forebody of model 57-B decreased the water resistance and the spray at speeds below the hump speed. Spray strips provided an effective means for the control of spray with the straight V sections used in the series but considerably increased the aerodynamic drag. Charts for the determination of the water resistance and the static properties of the model with 25 degrees dead rise and for the aerodynamic drag of all the models are included for use in design.
NASA Technical Reports Server (NTRS)
Mineck, R. E.
1977-01-01
Tests were conducted in the Langley V/STOL tunnel to determine the effect of the main-rotor wake on the aerodynamic characteristics of the rotor systems research aircraft. A 1/6-scale model with a 4-blade articulated rotor was used to determine the effect of the rotor wake for the compound configuration. Data were obtained over a range of angles of attack, angles of sideslip, auxiliary engine thrusts, rotor collective pitch angles, and rotor tip-path plane angles for several main-rotor advance ratios. Separate results are presented for the forces and moments on the airframe, the wing, and the tail. An analysis of the test data indicates significant changes in the aerodynamic characteristics. The rotor wake increases the longitudinal static stability, the effective dihedral, and the lateral static stability of the airframe. The rotor induces a downwash on the wing. This downwash decreases the wing lift and increases the drag. The asymmetrical rotor wake induces a differential lift across the wing and a subsequent rolling moment. These rotor induced effects on the wing become smaller with increasing forward speed.
Size effects on insect hovering aerodynamics: an integrated computational study.
Liu, H; Aono, H
2009-03-01
Hovering is a miracle of insects that is observed for all sizes of flying insects. Sizing effect in insect hovering on flapping-wing aerodynamics is of interest to both the micro-air-vehicle (MAV) community and also of importance to comparative morphologists. In this study, we present an integrated computational study of such size effects on insect hovering aerodynamics, which is performed using a biology-inspired dynamic flight simulator that integrates the modelling of realistic wing-body morphology, the modelling of flapping-wing and body kinematics and an in-house Navier-Stokes solver. Results of four typical insect hovering flights including a hawkmoth, a honeybee, a fruit fly and a thrips, over a wide range of Reynolds numbers from O(10(4)) to O(10(1)) are presented, which demonstrate the feasibility of the present integrated computational methods in quantitatively modelling and evaluating the unsteady aerodynamics in insect flapping flight. Our results based on realistically modelling of insect hovering therefore offer an integrated understanding of the near-field vortex dynamics, the far-field wake and downwash structures, and their correlation with the force production in terms of sizing and Reynolds number as well as wing kinematics. Our results not only give an integrated interpretation on the similarity and discrepancy of the near- and far-field vortex structures in insect hovering but also demonstrate that our methods can be an effective tool in the MAVs design.
Numerical simulation of the tip aerodynamics and acoustics test
NASA Astrophysics Data System (ADS)
Tejero E, F.; Doerffer, P.; Szulc, O.; Cross, J. L.
2016-04-01
The application of an efficient flow control system on helicopter rotor blades may lead to improved aerodynamic performance. Recently, our invention of Rod Vortex Generators (RVGs) has been analyzed for helicopter rotor blades in hover with success. As a step forward, the study has been extended to forward flight conditions. For this reason, a validation of the numerical modelling for a reference helicopter rotor (without flow control) is needed. The article presents a study of the flow-field of the AH-1G helicopter rotor in low-, medium- and high-speed forward flight. The CFD code FLOWer from DLR has proven to be a suitable tool for the aerodynamic analysis of the two-bladed rotor without any artificial wake modelling. It solves the URANS equations with LEA (Linear Explicit Algebraic stress) k-ω model using the chimera overlapping grids technique. Validation of the numerical model uses comparison with the detailed flight test data gathered by Cross J. L. and Watts M. E. during the Tip Aerodynamics and Acoustics Test (TAAT) conducted at NASA in 1981. Satisfactory agreements for all speed regimes and a presence of significant flow separation in high-speed forward flight suggest a possible benefit from the future implementation of RVGs. The numerical results based on the URANS approach are presented not only for a popular, low-speed case commonly used in rotorcraft community for CFD codes validation but preferably for medium- and high-speed test conditions that have not been published to date.
Coupled Aerodynamic-Thermal-Structural (CATS) Analysis
NASA Technical Reports Server (NTRS)
1995-01-01
Coupled Aerodynamic-Thermal-Structural (CATS) Analysis is a focused effort within the Numerical Propulsion System Simulation (NPSS) program to streamline multidisciplinary analysis of aeropropulsion components and assemblies. Multidisciplinary analysis of axial-flow compressor performance has been selected for the initial focus of this project. CATS will permit more accurate compressor system analysis by enabling users to include thermal and mechanical effects as an integral part of the aerodynamic analysis of the compressor primary flowpath. Thus, critical details, such as the variation of blade tip clearances and the deformation of the flowpath geometry, can be more accurately modeled and included in the aerodynamic analyses. The benefits of this coupled analysis capability are (1) performance and stall line predictions are improved by the inclusion of tip clearances and hot geometries, (2) design alternatives can be readily analyzed, and (3) higher fidelity analysis by researchers in various disciplines is possible. The goals for this project are a 10-percent improvement in stall margin predictions and a 2:1 speed-up in multidisciplinary analysis times. Working cooperatively with Pratt & Whitney, the Lewis CATS team defined the engineering processes and identified the software products necessary for streamlining these processes. The basic approach is to integrate the aerodynamic, thermal, and structural computational analyses by using data management and Non-Uniform Rational B-Splines (NURBS) based data mapping. Five software products have been defined for this task: (1) a primary flowpath data mapper, (2) a two-dimensional data mapper, (3) a database interface, (4) a blade structural pre- and post-processor, and (5) a computational fluid dynamics code for aerothermal analysis of the drum rotor. Thus far (1) a cooperative agreement has been established with Pratt & Whitney, (2) a Primary Flowpath Data Mapper has been prototyped and delivered to General Electric
X-34 Vehicle Aerodynamic Characteristics
NASA Technical Reports Server (NTRS)
Brauckmann, Gregory J.
1998-01-01
The X-34, being designed and built by the Orbital Sciences Corporation, is an unmanned sub-orbital vehicle designed to be used as a flying test bed to demonstrate key vehicle and operational technologies applicable to future reusable launch vehicles. The X-34 will be air-launched from an L-1011 carrier aircraft at approximately Mach 0.7 and 38,000 feet altitude, where an onboard engine will accelerate the vehicle to speeds above Mach 7 and altitudes to 250,000 feet. An unpowered entry will follow, including an autonomous landing. The X-34 will demonstrate the ability to fly through inclement weather, land horizontally at a designated site, and have a rapid turn-around capability. A series of wind tunnel tests on scaled models was conducted in four facilities at the NASA Langley Research Center to determine the aerodynamic characteristics of the X-34. Analysis of these test results revealed that longitudinal trim could be achieved throughout the design trajectory. The maximum elevon deflection required to trim was only half of that available, leaving a margin for gust alleviation and aerodynamic coefficient uncertainty. Directional control can be achieved aerodynamically except at combined high Mach numbers and high angles of attack, where reaction control jets must be used. The X-34 landing speed, between 184 and 206 knots, is within the capabilities of the gear and tires, and the vehicle has sufficient rudder authority to control the required 30-knot crosswind.
NASA Technical Reports Server (NTRS)
Ramsey, P.; Robertson, M. K.
1973-01-01
A test of a 0.004-scale MCR 0074 Baseline Launch Configuration Space Shuttle model was conducted in the NASA-MSFC 14 x 14-inch Trisonic Wind Tunnel (MSFC TWT 566). The objective of the test was to determine the effects of model parametric variations on aerodynamic static stability characteristics over a Mach number range from 0.6 to 4.96. Angles-of-attack from minus 10 deg to plus 10 deg at 0 deg sideslip and angles-of-sideslip from minus 10 deg to plus 10 deg at minus 5 deg, 0 deg, and plus 5 deg angle-of-attack were investigated. The basic configuration investigated was the integrated vehicle consisting of the orbiter, and external tank, and two solid rocket boosters. It was designated 03T9S3.
Numerical aerodynamic simulation facility feasibility study
NASA Technical Reports Server (NTRS)
1979-01-01
There were three major issues examined in the feasibility study. First, the ability of the proposed system architecture to support the anticipated workload was evaluated. Second, the throughput of the computational engine (the flow model processor) was studied using real application programs. Third, the availability reliability, and maintainability of the system were modeled. The evaluations were based on the baseline systems. The results show that the implementation of the Numerical Aerodynamic Simulation Facility, in the form considered, would indeed be a feasible project with an acceptable level of risk. The technology required (both hardware and software) either already exists or, in the case of a few parts, is expected to be announced this year. Facets of the work described include the hardware configuration, software, user language, and fault tolerance.
NASA Technical Reports Server (NTRS)
Peterson, Victor L.; Menees, Gene P.
1961-01-01
Tabulated results of a wind-tunnel investigation of the aerodynamic loads on a canard airplane model with a single vertical tail are presented for Mach numbers from 0.70 to 2.22. The Reynolds number for the measurements was 2.9 x 10(exp 6) based on the wing mean aerodynamic chord. The results include local static pressure coefficients measured on the wing, body, and vertical tail for angles of attack from -4 deg to + 16 deg, angles of sideslip of 0 deg and 5.3 deg, vertical-tail settings of 0 deg and 5 deg, and nominal canard deflections of 0 deg and 10 deg. Also included are section force and moment coefficients obtained from integrations of the local pressures and model-component force and moment coefficients obtained from integrations of the section coefficients. Geometric details of the model and the locations of the pressure orifices are shown. An index to the data contained herein is presented and definitions of nomenclature are given.
Aerodynamic analysis of hypersonic waverider aircraft
NASA Technical Reports Server (NTRS)
Sandlin, Doral R.; Pessin, David N.
1993-01-01
The purpose of this study is to validate two existing codes used by the Systems Analysis Branch at NASA ARC, and to modify the codes so they can be used to generate and analyze waverider aircraft at on-design and off-design conditions. To generate waverider configurations and perform the on-design analysis, the appropriately named Waverider code is used. The Waverider code is based on the Taylor-Maccoll equations. Validation is accomplished via a comparison with previously published results. The Waverider code is modified to incorporate a fairing to close off the base area of the waverider configuration. This creates a more realistic waverider. The Hypersonic Aircraft Vehicle Optimization Code (HAVOC) is used to perform the off-design analysis of waverider configurations generated by the Waverider code. Various approximate analysis methods are used by HAVOC to predict the aerodynamic characteristics, which are validated via a comparison with experimental results from a hypersonic test model.
Aerodynamic Simulation of Ice Accretion on Airfoils
NASA Technical Reports Server (NTRS)
Broeren, Andy P.; Addy, Harold E., Jr.; Bragg, Michael B.; Busch, Greg T.; Montreuil, Emmanuel
2011-01-01
This report describes recent improvements in aerodynamic scaling and simulation of ice accretion on airfoils. Ice accretions were classified into four types on the basis of aerodynamic effects: roughness, horn, streamwise, and spanwise ridge. The NASA Icing Research Tunnel (IRT) was used to generate ice accretions within these four types using both subscale and full-scale models. Large-scale, pressurized windtunnel testing was performed using a 72-in.- (1.83-m-) chord, NACA 23012 airfoil model with high-fidelity, three-dimensional castings of the IRT ice accretions. Performance data were recorded over Reynolds numbers from 4.5 x 10(exp 6) to 15.9 x 10(exp 6) and Mach numbers from 0.10 to 0.28. Lower fidelity ice-accretion simulation methods were developed and tested on an 18-in.- (0.46-m-) chord NACA 23012 airfoil model in a small-scale wind tunnel at a lower Reynolds number. The aerodynamic accuracy of the lower fidelity, subscale ice simulations was validated against the full-scale results for a factor of 4 reduction in model scale and a factor of 8 reduction in Reynolds number. This research has defined the level of geometric fidelity required for artificial ice shapes to yield aerodynamic performance results to within a known level of uncertainty and has culminated in a proposed methodology for subscale iced-airfoil aerodynamic simulation.
NASA Technical Reports Server (NTRS)
Newsom, William A., Jr.
1960-01-01
An investigation has been made to study the effect of ground proximity on the aerodynamic characteristics of two jet vertical-take-off-and-landing airplane models in which the fuselage remains in a horizontal attitude for the take-off and landing. The first model (called the tilt-wing model) had a tilting wing-engine assembly which was set at 90 deg incidence for the take-off and landing. The second model, called the deflected-jet model) had a cascade of retractable turning vanes to deflect the exhaust of the horizontally mounted jet engines downward for vertical take-off and landing while the entire model remained in a horizontal attitude. With the models at various heights above the ground in the take-off and landing configuration, the lift, drag, and pitching moment were measured and tuft surveys were made to determine the flow field caused by the jet exhaust. The tilt-wing model experienced a loss of lift of less than 3 percent near the ground. The deflected-jet model, however, suffered losses in lift as high as 45 percent near the ground because of a low pressure region under the model caused by the entrainment of air by the jet exhaust as it spread out along the ground. This loss in lift for the deflected-jet configuration could probably be reduced to less than 5 percent by the use of a longer landing gear and a high wing location.
Unsteady Aerodynamic Effects on the Flight Characteristics of an F-16XL Configuration
NASA Technical Reports Server (NTRS)
Wang, Zhongjun; Lan, C. Edward; Brandon, Jay M.
2000-01-01
Unsteady aerodynamic models based on windtunnel forced oscillation test data and analyzed with a fuzzy logic algorithm arc incorporated into an F-16XL flight simulation code. The reduced frequency needed in the unsteady models is numerically calculated by using a limited prior time history of state variables in a least-square sense. Numerical examples arc presented to show the accuracy of the calculated reduced frequency. Oscillatory control inputs are employed to demonstrate the differences in the flight characteristics based on unsteady and quasi-steady aerodynamic models. Application of the unsteady aerodynamic models is also presented and the results are compared with one set of F16XIL longitudinal maneuver flight data. It is shown that the main differences in dynamic response are in the lateral-directional characteristics, with the quasi-steady model being more stable than the flight vehicle, while the unsteady model being more unstable. Similar conclusions can also be made in a simulated rapid sideslipping roll. To improve unsteady aerodynamic modeling, it is recommended to acquire test data with coupled motions in pitch, roll and yaw.
Fundamental Aspects of the Aerodynamics of Turbojet Engine Combustors
NASA Technical Reports Server (NTRS)
Barrere, M.
1978-01-01
Aerodynamic considerations in the design of high performance combustors for turbojet engines are discussed. Aerodynamic problems concerning the preparation of the fuel-air mixture, the recirculation zone where primary combustion occurs, the secondary combustion zone, and the dilution zone were examined. An aerodynamic analysis of the entire primary chamber ensemble was carried out to determine the pressure drop between entry and exit. The aerodynamics of afterburn chambers are discussed. A model which can be used to investigate the evolution of temperature, pressure, and rate and efficiency of combustion the length of the chamber was developed.
NASA Technical Reports Server (NTRS)
Mercer, C. E.; Carson, G. T., Jr.
1979-01-01
The influence of upper-surface nacelle exhaust flow on the aerodynamic characteristics of a supersonic cruise aircraft research configuration was investigated in a 16 foot transonic tunnel over a range of Mach numbers from 0.60 to 1.20. The arrow-wing transport configuration with engines suspended over the wing was tested at angles of attack from -4 deg to 6 deg and jet total pressure ratios from 1 to approximately 13. Wing-tip leading edge flap deflections of -10 deg to 10 deg were tested with the wing-body configuration. Various nacelle locations (chordwise, spanwise, and vertical) were tested over the ranges of Mach numbers, angles of attack, and jet total-pressure ratios. The results show that reflecting the wing-tip leading edge flap from 0 deg to -10 deg increased the maximum lift-drag ratio by 1.0 at subsonic speeds. Jet exhaust interference effects were negligible.
NASA Technical Reports Server (NTRS)
Parkinson, John B; Olson, Roland E; Draley, Eugene C; Luoma, Arvo A
1943-01-01
A series of related forms of flying-boat hulls representing various degrees of compromise between aerodynamic and hydrodynamic requirements was tested in Langley Tank No. 1 and in the Langley 8-foot high-speed tunnel. The purpose of the investigation was to provide information regarding the penalties in water performance resulting from further aerodynamic refinement and, as a corollary, to provide information regarding the penalties in range or payload resulting from the retention of certain desirable hydrodynamic characteristics. The information should form a basis for over-all improvements in hull form.
NASA Technical Reports Server (NTRS)
Tweedt, Daniel L.
2014-01-01
Computational Aerodynamic simulations of a 1484 ft/sec tip speed quiet high-speed fan system were performed at five different operating points on the fan operating line, in order to provide detailed internal flow field information for use with fan acoustic prediction methods presently being developed, assessed and validated. The fan system is a sub-scale, low-noise research fan/nacelle model that has undergone experimental testing in the 9- by 15-foot Low Speed Wind Tunnel at the NASA Glenn Research Center. Details of the fan geometry, the computational fluid dynamics methods, the computational grids, and various computational parameters relevant to the numerical simulations are discussed. Flow field results for three of the five operating points simulated are presented in order to provide a representative look at the computed solutions. Each of the five fan aerodynamic simulations involved the entire fan system, which includes a core duct and a bypass duct that merge upstream of the fan system nozzle. As a result, only fan rotational speed and the system bypass ratio, set by means of a translating nozzle plug, were adjusted in order to set the fan operating point, leading to operating points that lie on a fan operating line and making mass flow rate a fully dependent parameter. The resulting mass flow rates are in good agreement with measurement values. Computed blade row flow fields at all fan operating points are, in general, aerodynamically healthy. Rotor blade and fan exit guide vane flow characteristics are good, including incidence and deviation angles, chordwise static pressure distributions, blade surface boundary layers, secondary flow structures, and blade wakes. Examination of the computed flow fields reveals no excessive or critical boundary layer separations or related secondary-flow problems, with the exception of the hub boundary layer at the core duct entrance. At that location a significant flow separation is present. The region of local flow
Photogrammetry of a Hypersonic Inflatable Aerodynamic Decelerator
NASA Technical Reports Server (NTRS)
Kushner, Laura Kathryn; Littell, Justin D.; Cassell, Alan M.
2013-01-01
In 2012, two large-scale models of a Hypersonic Inflatable Aerodynamic decelerator were tested in the National Full-Scale Aerodynamic Complex at NASA Ames Research Center. One of the objectives of this test was to measure model deflections under aerodynamic loading that approximated expected flight conditions. The measurements were acquired using stereo photogrammetry. Four pairs of stereo cameras were mounted inside the NFAC test section, each imaging a particular section of the HIAD. The views were then stitched together post-test to create a surface deformation profile. The data from the photogram- metry system will largely be used for comparisons to and refinement of Fluid Structure Interaction models. This paper describes how a commercial photogrammetry system was adapted to make the measurements and presents some preliminary results.
Powered-Lift Aerodynamics and Acoustics. [conferences
NASA Technical Reports Server (NTRS)
1976-01-01
Powered lift technology is reviewed. Topics covered include: (1) high lift aerodynamics; (2) high speed and cruise aerodynamics; (3) acoustics; (4) propulsion aerodynamics and acoustics; (5) aerodynamic and acoustic loads; and (6) full-scale and flight research.
Aerodynamic Design Study of an Advanced Active Twist Rotor
NASA Technical Reports Server (NTRS)
Sekula, Martin K.; Wilbur, Matthew L.; Yeager, William T., Jr.
2003-01-01
An Advanced Active Twist Rotor (AATR) is currently being developed by the U.S. Army Vehicle Technology Directorate at NASA Langley Research Center. As a part of this effort, an analytical study was conducted to determine the impact of blade geometry on active-twist performance and, based on those findings, propose a candidate aerodynamic design for the AATR. The process began by creating a baseline design which combined the dynamic design of the original Active Twist Rotor and the aerodynamic design of a high lift rotor concept. The baseline model was used to conduct a series of parametric studies to examine the effect of linear blade twist and blade tip sweep, droop, and taper on active-twist performance. Rotor power requirements and hub vibration were also examined at flight conditions ranging from hover to advance ratio = 0.40. A total of 108 candidate designs were analyzed using the second-generation version of the Comprehensive Analytical Model of Rotorcraft Aerodynamics and Dynamics (CAMRAD II) code. The study concluded that the vibration reduction capabilities of a rotor utilizing controlled, strain-induced twisting are enhanced through the incorporation of blade tip sweep, droop, and taper into the blade design, while they are degraded by increasing the nose-down linear blade twist. Based on the analysis of rotor power, hub vibration, and active-twist response, a candidate aerodynamic design for the AATR consisting of a blade with approximately 10 degrees of linear blade twist and a blade tip design with 30 degree sweep, 10 degree droop, and 2.5:1 taper ratio over the outer five percent of the blade is proposed.
NASA Astrophysics Data System (ADS)
Isaev, S. A.; Baranov, P. A.; Sudakov, A. G.; Ermakov, A. M.
2015-01-01
The Reynolds equations closed using the Menter shear-stress-transfer model modified with allowance for the curvature of flow lines have been numerically solved using multiblock computational technologies. The obtained solution has been used to analyze subsonic flow past a thick (37.5% chord) airfoil with slot suction in circular vortex cells intended for the Ecology and Progress (Ekologiya i Progress, EKIP) aircraft project in comparison to a distributed (from the central body surface) suction at fixed values of the total volume flow rate (0.02121) and Reynolds number (105) in a range of Mach numbers from 0 to 0.4. This analysis revealed a significant (up to tenfold) decrease in the bow drag (determined with allowance for the energy losses) and a large (by an order of magnitude) increase in the aerodynamic efficiency of the thick airfoil containing vortex cells with slot suction, which reached up to 160.
NASA Technical Reports Server (NTRS)
Stimpert, D. L.
1978-01-01
An acoustic and aerodynamic test program was conducted on a 1/6.25 scale model of the Quiet, Clean, Short-Haul Experimental Engine (QCSEE) forward thrust over-the-wing (OTW) nozzle and OTW thrust reverser. In reverse thrust, the effect of reverser geometry was studied by parametric variations in blocker spacing, blocker height, lip angle, and lip length. Forward thrust nozzle tests determined the jet noise levels of the cruise and takeoff nozzles, the effect of opening side doors to achieve takeoff thrust, and scrubbing noise of the cruise and takeoff jet on a simulated wing surface. Velocity profiles are presented for both forward and reverse thrust nozzles. An estimate of the reverse thrust was made utilizing the measured centerline turning angle.
NASA Astrophysics Data System (ADS)
Varshney, Kapil; Chang, Song; Wang, Z. Jane
2013-05-01
Falling parallelograms exhibit coupled motion of autogyration and tumbling, similar to the motion of falling tulip seeds, unlike maple seeds which autogyrate but do not tumble, or rectangular cards which tumble but do not gyrate. This coupled tumbling and autogyrating motion are robust, when card parameters, such as aspect ratio, internal angle, and mass density, are varied. We measure the three-dimensional (3D) falling kinematics of the parallelograms and quantify their descending speed, azimuthal rotation, tumbling rotation, and cone angle in each falling. The cone angle is insensitive to the variation of the card parameters, and the card tumbling axis does not overlap with but is close to the diagonal axis. In addition to this connection to the dynamics of falling seeds, these trajectories provide an ideal set of data to analyze 3D aerodynamic force and torque at an intermediate range of Reynolds numbers, and the results will be useful for constructing 3D aerodynamic force and torque models. Tracking these free falling trajectories gives us a nonintrusive method for deducing instantaneous aerodynamic forces. We determine the 3D aerodynamic forces and torques based on Newton-Euler equations. The dynamical analysis reveals that, although the angle of attack changes dramatically during tumbling, the aerodynamic forces have a weak dependence on the angle of attack. The aerodynamic lift is dominated by the coupling of translational and rotational velocities. The aerodynamic torque has an unexpectedly large component perpendicular to the card. The analysis of the Euler equation suggests that this large torque is related to the deviation of the tumbling axis from the principle axis of the card.
NASA Astrophysics Data System (ADS)
Rege, Alok Ashok
Insect flight comes with a lot of intricacies that cannot be explained by conventional aerodynamics. Even with their small-size, insects have the ability to generate the required aerodynamic forces using high frequency flapping motion of their wings to perform different maneuvers. The maneuverability obtained by these flyers using flapping motion belies the classical aerodynamics theory and calls for a new approach to study this highly unsteady aerodynamics. Research is on to find new ways to realize the flight capabilities of these insects and engineer a micro-flyer which would have various applications, ranging from autonomous pollination of crop fields and oil & gas exploration to area surveillance and detection & rescue missions. In this research, a parametric study of flapping trajectories is performed using a two-dimensional wing to identify the factors that affect the force production. These factors are then non-dimensionalized and used in a design of experiments set-up to conduct sensitivity analysis. A procedure to determine an aerodynamic model comprising cycle-averaged force coefficients is described. This aerodynamic model is then used in a nonlinear dynamics framework to perform flight dynamics analysis using a micro-flyer with model properties based on Drosophila. Stability analysis is conducted to determine different steady state flight conditions that could achieved by the micro-flyer with the given model properties. The effect of scaling the mass properties is discussed. An LQR design is used for closed-loop control. Open and closed-loop simulations are performed. The results show that nonlinear dynamics framework can be used to determine values for model properties of a micro-flyer that would enable it to perform different flight maneuvers.
Applied computational aerodynamics
Henne, P.A.
1990-01-01
The present volume discusses the original development of the panel method, the mapping solutions and singularity distributions of linear potential schemes, the capabilities of full-potential, Euler, and Navier-Stokes schemes, the use of the grid-generation methodology in applied aerodynamics, subsonic airfoil design, inverse airfoil design for transonic applications, the divergent trailing-edge airfoil innovation in CFD, Euler and potential computational results for selected aerodynamic configurations, and the application of CFD to wing high-lift systems. Also discussed are high-lift wing modifications for an advanced-capability EA-6B aircraft, Navier-Stokes methods for internal and integrated propulsion system flow predictions, the use of zonal techniques for analysis of rotor-stator interaction, CFD applications to complex configurations, CFD applications in component aerodynamic design of the V-22, Navier-Stokes computations of a complete F-16, CFD at supersonic/hypersonic speeds, and future CFD developments.
In vivo recording of aerodynamic force with an aerodynamic force platform: from drones to birds.
Lentink, David; Haselsteiner, Andreas F; Ingersoll, Rivers
2015-03-06
Flapping wings enable flying animals and biomimetic robots to generate elevated aerodynamic forces. Measurements that demonstrate this capability are based on experiments with tethered robots and animals, and indirect force calculations based on measured kinematics or airflow during free flight. Remarkably, there exists no method to measure these forces directly during free flight. Such in vivo recordings in freely behaving animals are essential to better understand the precise aerodynamic function of their flapping wings, in particular during the downstroke versus upstroke. Here, we demonstrate a new aerodynamic force platform (AFP) for non-intrusive aerodynamic force measurement in freely flying animals and robots. The platform encloses the animal or object that generates fluid force with a physical control surface, which mechanically integrates the net aerodynamic force that is transferred to the earth. Using a straightforward analytical solution of the Navier-Stokes equation, we verified that the method is accurate. We subsequently validated the method with a quadcopter that is suspended in the AFP and generates unsteady thrust profiles. These independent measurements confirm that the AFP is indeed accurate. We demonstrate the effectiveness of the AFP by studying aerodynamic weight support of a freely flying bird in vivo. These measurements confirm earlier findings based on kinematics and flow measurements, which suggest that the avian downstroke, not the upstroke, is primarily responsible for body weight support during take-off and landing.
In vivo recording of aerodynamic force with an aerodynamic force platform: from drones to birds
Lentink, David; Haselsteiner, Andreas F.; Ingersoll, Rivers
2015-01-01
Flapping wings enable flying animals and biomimetic robots to generate elevated aerodynamic forces. Measurements that demonstrate this capability are based on experiments with tethered robots and animals, and indirect force calculations based on measured kinematics or airflow during free flight. Remarkably, there exists no method to measure these forces directly during free flight. Such in vivo recordings in freely behaving animals are essential to better understand the precise aerodynamic function of their flapping wings, in particular during the downstroke versus upstroke. Here, we demonstrate a new aerodynamic force platform (AFP) for non-intrusive aerodynamic force measurement in freely flying animals and robots. The platform encloses the animal or object that generates fluid force with a physical control surface, which mechanically integrates the net aerodynamic force that is transferred to the earth. Using a straightforward analytical solution of the Navier–Stokes equation, we verified that the method is accurate. We subsequently validated the method with a quadcopter that is suspended in the AFP and generates unsteady thrust profiles. These independent measurements confirm that the AFP is indeed accurate. We demonstrate the effectiveness of the AFP by studying aerodynamic weight support of a freely flying bird in vivo. These measurements confirm earlier findings based on kinematics and flow measurements, which suggest that the avian downstroke, not the upstroke, is primarily responsible for body weight support during take-off and landing. PMID:25589565
Analysis of aerodynamic noise generated from inclined circular cylinder
NASA Astrophysics Data System (ADS)
Haramoto, Yasutake; Yasuda, Shouji; Matsuzaki, Kazuyoshi; Munekata, Mizue; Ohba, Hideki
2000-06-01
Making clear the generation mechanism of fluid dynamic noise is essential to reduce noise deriving from turbomachinery. The analysis of the aerodynamic noise generated from circular cylinder is carried out numerically and experimentally in a low noise wind tunnel. In this study, aerodynamic sound radiated from a circular cylinder in uniform flow is predicted numerically by the following two step method. First, the three-dimensional unsteady incompressible Navier-Stokes equation is solved using the high order accurate upwind scheme. Next, the sound pressure level at the observed point is calculated from the fluctuating surface pressure on the cylinder, based on modified Lighthill-Curl’s equation. It is worth to note that the noise generated from the model is reduced rapidly when it is inclined against the mean flow. In other words, the peak level of the radiated noise decreases rapidly with inclination of the circular cylinder. The simulated SPL for the inclined circular cylinder is compared with the measured value, and good agreement is obtained for the peak spectrum frequency of the sound pressure level and tendency of noise reduction. So we expect that the change of flow structures makes reduction of the aerodynamic noise from the inclined models.
Doubrawa, P.; Barthelmie, R. J.; Wang, H.; ...
2016-10-03
The contribution of wake meandering and shape asymmetry to load and power estimates is quantified by comparing aeroelastic simulations initialized with different inflow conditions: an axisymmetric base wake, an unsteady stochastic shape wake, and a large-eddy simulation with rotating actuator-line turbine representation. Time series of blade-root and tower base bending moments are analyzed. We find that meandering has a large contribution to the fluctuation of the loads. Moreover, considering the wake edge intermittence via the stochastic shape model improves the simulation of load and power fluctuations and of the fatigue damage equivalent loads. Furthermore, these results indicate that the stochasticmore » shape wake simulator is a valuable addition to simplified wake models when seeking to obtain higher-fidelity computationally inexpensive predictions of loads and power.« less
Doubrawa, P.; Barthelmie, R. J.; Wang, H.; Churchfield, M. J.
2016-10-03
The contribution of wake meandering and shape asymmetry to load and power estimates is quantified by comparing aeroelastic simulations initialized with different inflow conditions: an axisymmetric base wake, an unsteady stochastic shape wake, and a large-eddy simulation with rotating actuator-line turbine representation. Time series of blade-root and tower base bending moments are analyzed. We find that meandering has a large contribution to the fluctuation of the loads. Moreover, considering the wake edge intermittence via the stochastic shape model improves the simulation of load and power fluctuations and of the fatigue damage equivalent loads. Furthermore, these results indicate that the stochastic shape wake simulator is a valuable addition to simplified wake models when seeking to obtain higher-fidelity computationally inexpensive predictions of loads and power.
Turbine disk cavity aerodynamics and heat transfer
NASA Technical Reports Server (NTRS)
Johnson, B. V.; Daniels, W. A.
1992-01-01
Experiments were conducted to define the nature of the aerodynamics and heat transfer for the flow within the disk cavities and blade attachments of a large-scale model, simulating the Space Shuttle Main Engine (SSME) turbopump drive turbines. These experiments of the aerodynamic driving mechanisms explored the following: (1) flow between the main gas path and the disk cavities; (2) coolant flow injected into the disk cavities; (3) coolant density; (4) leakage flows through the seal between blades; and (5) the role that each of these various flows has in determining the adiabatic recovery temperature at all of the critical locations within the cavities. The model and the test apparatus provide close geometrical and aerodynamic simulation of all the two-stage cavity flow regions for the SSME High Pressure Fuel Turbopump and the ability to simulate the sources and sinks for each cavity flow.
Parameter identification for nonlinear aerodynamic systems
NASA Technical Reports Server (NTRS)
Pearson, Allan E.
1990-01-01
Parameter identification for nonlinear aerodynamic systems is examined. It is presumed that the underlying model can be arranged into an input/output (I/O) differential operator equation of a generic form. The algorithm estimation is especially efficient since the equation error can be integrated exactly given any I/O pair to obtain an algebraic function of the parameters. The algorithm for parameter identification was extended to the order determination problem for linear differential system. The degeneracy in a least squares estimate caused by feedback was addressed. A method of frequency analysis for determining the transfer function G(j omega) from transient I/O data was formulated using complex valued Fourier based modulating functions in contrast with the trigonometric modulating functions for the parameter estimation problem. A simulation result of applying the algorithm is given under noise-free conditions for a system with a low pass transfer function.
Aerodynamic Shape Optimization using an Evolutionary Algorithm
NASA Technical Reports Server (NTRS)
Hoist, Terry L.; Pulliam, Thomas H.
2003-01-01
A method for aerodynamic shape optimization based on an evolutionary algorithm approach is presented and demonstrated. Results are presented for a number of model problems to access the effect of algorithm parameters on convergence efficiency and reliability. A transonic viscous airfoil optimization problem-both single and two-objective variations is used as the basis for a preliminary comparison with an adjoint-gradient optimizer. The evolutionary algorithm is coupled with a transonic full potential flow solver and is used to optimize the inviscid flow about transonic wings including multi-objective and multi-discipline solutions that lead to the generation of pareto fronts. The results indicate that the evolutionary algorithm approach is easy to implement, flexible in application and extremely reliable.
Aerodynamic Shape Optimization using an Evolutionary Algorithm
NASA Technical Reports Server (NTRS)
Holst, Terry L.; Pulliam, Thomas H.; Kwak, Dochan (Technical Monitor)
2003-01-01
A method for aerodynamic shape optimization based on an evolutionary algorithm approach is presented and demonstrated. Results are presented for a number of model problems to access the effect of algorithm parameters on convergence efficiency and reliability. A transonic viscous airfoil optimization problem, both single and two-objective variations, is used as the basis for a preliminary comparison with an adjoint-gradient optimizer. The evolutionary algorithm is coupled with a transonic full potential flow solver and is used to optimize the inviscid flow about transonic wings including multi-objective and multi-discipline solutions that lead to the generation of pareto fronts. The results indicate that the evolutionary algorithm approach is easy to implement, flexible in application and extremely reliable.
Experimental investigation of the aerodynamic characteristics for a winged-cone concept
NASA Technical Reports Server (NTRS)
Phillips, W. Pelham; Brauckmann, Gregory J.; Micol, John R.; Woods, William C.
1987-01-01
Experimental longitudinal and lateral-directional aerodynamics were obtained for a generic aerodynamics were obtaiend for a generic winged-cone configuration having possible application as a transatmospheric vehicle concept. Data were obtained at Mach numbers from 0.6 to 20.0; Reynolds numbers, based on model length, between 2.5 and 5.3 million; and angles of attack from -4 to 20 deg. Results indicate a longitudinal center-of-pressure travel of about 23 percent of the fuselage length for the test Mach number range, with longitudinal instabilities noted at high-supersonic to hypersonic Mach numbers. These instabilities are coupled with directional instability at similar Mach numbers. Predictions with analytic codes, namely, the USAF DATCOM and the tangent-cone option of the Hypersonic Arbitrary Body Program, provided fair agreement with the experimental aerodynamic characteristics at low angles-of-attack.
Training Data Requirement for a Neural Network to Predict Aerodynamic Coefficients
NASA Technical Reports Server (NTRS)
Korsmeyer, David (Technical Monitor); Rajkumar, T.; Bardina, Jorge
2003-01-01
Basic aerodynamic coefficients are modeled as functions of angle of attack, speed brake deflection angle, Mach number, and side slip angle. Most of the aerodynamic parameters can be well-fitted using polynomial functions. We previously demonstrated that a neural network is a fast, reliable way of predicting aerodynamic coefficients. We encountered few under fitted and/or over fitted results during prediction. The training data for the neural network are derived from wind tunnel test measurements and numerical simulations. The basic questions that arise are: how many training data points are required to produce an efficient neural network prediction, and which type of transfer functions should be used between the input-hidden layer and hidden-output layer. In this paper, a comparative study of the efficiency of neural network prediction based on different transfer functions and training dataset sizes is presented. The results of the neural network prediction reflect the sensitivity of the architecture, transfer functions, and training dataset size.
Aerodynamic and aerothermodynamic trade-off analysis of a small hypersonic flying test bed
NASA Astrophysics Data System (ADS)
Pezzella, Giuseppe
2011-08-01
This paper deals with the aerodynamic and aerothermodynamic trade-off analysis aiming to design a small hypersonic flying test bed with a relatively simple vehicle architecture. Such vehicle will have to be launched with a sounding rocket and shall re-enter the Earth atmosphere allowing to perform several experiments on critical re-entry technologies such as boundary-layer transition and shock-shock interaction phenomena. The flight shall be conducted at hypersonic Mach number, in the range 6-8 at moderate angles of attack. In the paper some design analyses are shown as, for example, the longitudinal and lateral-directional stability analysis. A preliminary optimization of the configuration has been also done to improve the aerodynamic performance and stability of the vehicle. Several design results, based both on engineering approach and computational fluid dynamics, are reported and discussed in the paper. The aerodynamic model of vehicle is also provided.
Study of aerodynamic technology for single-cruise-engine V/STOL fighter/attack aircraft
NASA Technical Reports Server (NTRS)
Hess, J. R.; Bear, R. L.
1982-01-01
A viable, single engine, supersonic V/STOL fighter/attack aircraft concept was defined. This vectored thrust, canard wing configuration utilizes an advanced technology separated flow engine with fan stream burning. The aerodynamic characteristics of this configuration were estimated and performance evaluated. Significant aerodynamic and aerodynamic propulsion interaction uncertainties requiring additional investigation were identified. A wind tunnel model concept and test program to resolve these uncertainties and validate the aerodynamic prediction methods were defined.
Bat flight generates complex aerodynamic tracks.
Hedenström, A; Johansson, L C; Wolf, M; von Busse, R; Winter, Y; Spedding, G R
2007-05-11
The flapping flight of animals generates an aerodynamic footprint as a time-varying vortex wake in which the rate of momentum change represents the aerodynamic force. We showed that the wakes of a small bat species differ from those of birds in some important respects. In our bats, each wing generated its own vortex loop. Also, at moderate and high flight speeds, the circulation on the outer (hand) wing and the arm wing differed in sign during the upstroke, resulting in negative lift on the hand wing and positive lift on the arm wing. Our interpretations of the unsteady aerodynamic performance and function of membranous-winged, flapping flight should change modeling strategies for the study of equivalent natural and engineered flying devices.
Aerodynamics of saccate pollen and its implications for wind pollination.
Schwendemann, Andrew B; Wang, George; Mertz, Meredith L; McWilliams, Ryan T; Thatcher, Scott L; Osborn, Jeffrey M
2007-08-01
Pollen grains of many wind-pollinated plants contain 1-3 air-filled bladders, or sacci. Sacci are thought to help orient the pollen grain in the pollination droplet. Sacci also increase surface area of the pollen grain, yet add minimal mass, thereby increasing dispersal distance; however, this aerodynamic hypothesis has not been tested in a published study. Using scanning electron and transmission electron microscopy, mathematical modeling, and the saccate pollen of three extant conifers with structurally different pollen grains (Pinus, Falcatifolium, Dacrydium), we developed a computational model to investigate pollen flight. The model calculates terminal settling velocity based on structural characters of the pollen grain, including lengths, widths, and depths of the main body and sacci; angle of saccus rotation; and thicknesses of the saccus wall, endoreticulations, intine, and exine. The settling speeds predicted by the model were empirically validated by stroboscopic photography. This study is the first to quantitatively demonstrate the adaptive significance of sacci for the aerodynamics of wind pollination. Modeling pollen both with and without sacci indicated that sacci can reduce pollen settling speeds, thereby increasing dispersal distance, with the exception of pollen grains having robust endoreticulations and those with thick saccus walls. Furthermore, because the mathematical model is based on structural characters and error propagation methods show that the model yields valid results when sample sizes are small, the flight dynamics of fossil pollen can be investigated. Several fossils were studied, including bisaccate (Pinus, Pteruchus, Caytonanthus), monosaccate (Gothania), and nonsaccate (Monoletes) pollen types.
Aerodynamic investigations of a disc-wing
NASA Astrophysics Data System (ADS)
Dumitrache, Alexandru; Frunzulica, Florin; Grigorescu, Sorin
2017-01-01
The purpose of this paper is to evaluate the aerodynamic characteristics of a wing-disc, for a civil application in the fire-fighting system. The aerodynamic analysis is performed using a CFD code, named ANSYS Fluent, in the flow speed range up to 25 m/s, at lower and higher angle of attack. The simulation is three-dimensional, using URANS completed by a SST turbulence model. The results are used to examine the flow around the disc with increasing angle of attack and the structure of the wake.
Unstructured mesh algorithms for aerodynamic calculations
NASA Technical Reports Server (NTRS)
Mavriplis, D. J.
1992-01-01
The use of unstructured mesh techniques for solving complex aerodynamic flows is discussed. The principle advantages of unstructured mesh strategies, as they relate to complex geometries, adaptive meshing capabilities, and parallel processing are emphasized. The various aspects required for the efficient and accurate solution of aerodynamic flows are addressed. These include mesh generation, mesh adaptivity, solution algorithms, convergence acceleration, and turbulence modeling. Computations of viscous turbulent two-dimensional flows and inviscid three-dimensional flows about complex configurations are demonstrated. Remaining obstacles and directions for future research are also outlined.
Transpiration effects in perforated plate aerodynamics
NASA Astrophysics Data System (ADS)
Szwaba, R.; Ochrymiuk, T.
2016-10-01
Perforated walls find a wide use as a method of flow control and effusive cooling. Experimental investigations of the gas flow past perforated plate with microholes (110μm) were carried out. The wide range of pressure at the inlet were investigated. Two distinguishable flow regimes were obtained: laminar and turbulent regime.The results are in good agreement with theory, simulations and experiments on large scale perforated plates and compressible flows in microtubules. Formulation of the transpiration law was associated with the porous plate aerodynamics properties. Using a model of transpiration flow the “aerodynamic porosity” could be determined for microholes.
Churchfield, M. J.; Michalakes, J.; Vanderwende, B.; Lee, S.; Sprague, M. A.; Lundquist, J. K.; Moriarty, P. J.
2013-10-01
Wind plant aerodynamics are directly affected by the microscale weather, which is directly influenced by the mesoscale weather. Microscale weather refers to processes that occur within the atmospheric boundary layer with the largest scales being a few hundred meters to a few kilometers depending on the atmospheric stability of the boundary layer. Mesoscale weather refers to large weather patterns, such as weather fronts, with the largest scales being hundreds of kilometers wide. Sometimes microscale simulations that capture mesoscale-driven variations (changes in wind speed and direction over time or across the spatial extent of a wind plant) are important in wind plant analysis. In this paper, we present our preliminary work in coupling a mesoscale weather model with a microscale atmospheric large-eddy simulation model. The coupling is one-way beginning with the weather model and ending with a computational fluid dynamics solver using the weather model in coarse large-eddy simulation mode as an intermediary. We simulate one hour of daytime moderately convective microscale development driven by the mesoscale data, which are applied as initial and boundary conditions to the microscale domain, at a site in Iowa. We analyze the time and distance necessary for the smallest resolvable microscales to develop.
NASA Technical Reports Server (NTRS)
Horvath, Thomas J.; OConnell, Tod F.; Cheatwood, F. McNeil; Prabhu, Ramadas K.; Alter, Stephen J.
2002-01-01
Aerodynamic wind-tunnel screening tests were conducted on a 0.029 scale model of a proposed Mars Surveyor 2001 Precision Lander (70 deg half angle spherically blunted cone with a conical afterbody). The primary experimental objective was to determine the effectiveness of a single flap to trim the vehicle at incidence during a lifting hypersonic planetary entry. The laminar force and moment data, presented in the form of coefficients, and shock patterns from schlieren photography were obtained in the NASA Langley Aerothermodynamic Laboratory for post-normal shock Reynolds numbers (based on forebody diameter) ranging from 2,637 to 92,350, angles of attack ranging from 0 tip to 23 degrees at 0 and 2 degree sideslip, and normal-shock density ratios of 5 and 12. Based upon the proposed entry trajectory of the 2001 Lander, the blunt body heavy gas tests in CF, simulate a Mach number of approximately 12 based upon a normal shock density ratio of 12 in flight at Mars. The results from this experimental study suggest that when traditional means of providing aerodynamic trim for this class of planetary entry vehicle are not possible (e.g. offset c.g.), a single flap can provide similar aerodynamic performance. An assessment of blunt body aerodynamic effects attributed to a real gas were obtained by synergistic testing in Mach 6 ideal-air at a comparable Reynolds number. From an aerodynamic perspective, an appropriately sized flap was found to provide sufficient trim capability at the desired L/D for precision landing. Inviscid hypersonic flow computations using an unstructured grid were made to provide a quick assessment of the Lander aerodynamics. Navier-Stokes computational predictions were found to be in very good agreement with experimental measurement.
Aerodynamic influences on atmospheric in situ measurements from sounding rockets
NASA Astrophysics Data System (ADS)
Gumbel, Jörg
2001-06-01
Sounding rockets are essential tools for studies of the mesosphere and lower thermosphere. However, in situ measurements from rockets are potentially subject to a number of perturbations related to the gas flow around the vehicle. This paper reviews the aerodynamic principles behind these perturbations. With respect to both data analysis and experiment design, there is a substantial need for improved understanding of aerodynamic effects. Any such analysis is complicated by the different flow regimes experienced during a rocket flight through the rarefied environment of the mesosphere and thermosphere. Numerical studies are presented using the Direct Simulation Monte Carlo (DSMC) approach, which is based on a tracing of individual molecules. Complementary experiments have been performed in a low-density wind tunnel. These experiments are crucial for the development of appropriate model parameterization. However, direct similarity between scaled wind tunnel results and arbitrary atmospheric flight conditions is usually difficult to achieve. Density, velocity, and temperature results are presented for different payload geometries and flow conditions. These illustrate a wide range of aerodynamic effects representative for rocket flights in the mesosphere and lower thermosphere.
NASA Technical Reports Server (NTRS)
Rajkumar, T.; Bardina, Jorge; Clancy, Daniel (Technical Monitor)
2002-01-01
Wind tunnels use scale models to characterize aerodynamic coefficients, Wind tunnel testing can be slow and costly due to high personnel overhead and intensive power utilization. Although manual curve fitting can be done, it is highly efficient to use a neural network to define the complex relationship between variables. Numerical simulation of complex vehicles on the wide range of conditions required for flight simulation requires static and dynamic data. Static data at low Mach numbers and angles of attack may be obtained with simpler Euler codes. Static data of stalled vehicles where zones of flow separation are usually present at higher angles of attack require Navier-Stokes simulations which are costly due to the large processing time required to attain convergence. Preliminary dynamic data may be obtained with simpler methods based on correlations and vortex methods; however, accurate prediction of the dynamic coefficients requires complex and costly numerical simulations. A reliable and fast method of predicting complex aerodynamic coefficients for flight simulation I'S presented using a neural network. The training data for the neural network are derived from numerical simulations and wind-tunnel experiments. The aerodynamic coefficients are modeled as functions of the flow characteristics and the control surfaces of the vehicle. The basic coefficients of lift, drag and pitching moment are expressed as functions of angles of attack and Mach number. The modeled and training aerodynamic coefficients show good agreement. This method shows excellent potential for rapid development of aerodynamic models for flight simulation. Genetic Algorithms (GA) are used to optimize a previously built Artificial Neural Network (ANN) that reliably predicts aerodynamic coefficients. Results indicate that the GA provided an efficient method of optimizing the ANN model to predict aerodynamic coefficients. The reliability of the ANN using the GA includes prediction of aerodynamic
Aerodynamic properties of turbulent combustion fields
NASA Technical Reports Server (NTRS)
Hsiao, C. C.; Oppenheim, A. K.
1985-01-01
Flow fields involving turbulent flames in premixed gases under a variety of conditions are modeled by the use of a numerical technique based on the random vortex method to solve the Navier-Stokes equations and a flame propagation algorithm to trace the motion of the front and implement the Huygens principle, both due to Chorin. A successive over-relaxation hybrid method is applied to solve the Euler equation for flows in an arbitrarily shaped domain. The method of images, conformal transformation, and the integral-equation technique are also used to treat flows in special cases, according to their particular requirements. Salient features of turbulent flame propagation in premixed gases are interpreted by relating them to the aerodynamic properties of the flow field. Included among them is the well-known cellular structure of flames stabilized by bluff bodies, as well as the formation of the characteristic tulip shape of flames propagating in ducts. In its rudimentary form, the mechanism of propagation of a turbulent flame is shown to consist of: (1) rotary motion of eddies at the flame front, (2) self-advancement of the front at an appropriate normal burning speed, and (3) dynamic effects of expansion due to exothermicity of the combustion reaction. An idealized model is used to illustrate these fundamental mechanisms and to investigate basic aerodynamic features of flames in premixed gases. The case of a confined flame stabilized behind a rearward-facing step is given particular care and attention. Solutions are shown to be in satisfactory agreement with experimental results, especially with respect to global properties such as the average velocity profiles and reattachment length.
Aerodynamics and Aeroelasticity Calculations of Flapping Motion for Micro Air Vehicle
2009-08-24
Nastran program which uses Doublet Lattice Method for aerodynamic analysis and couples with its structural model . The program could handle also gust...numerical calculations a. 3D flat plate methods could be calculated with MSC. Nastran aeroelastic module. Conventional MSC.Nastran Structural Model performs...platforms will modelled by Direct Numerical Simulation technique (DNS) using a computer programme, based on an available commercial code Star-CD (or
Airfoil Ice-Accretion Aerodynamics Simulation
NASA Technical Reports Server (NTRS)
Bragg, Michael B.; Broeren, Andy P.; Addy, Harold E.; Potapczuk, Mark G.; Guffond, Didier; Montreuil, E.
2007-01-01
NASA Glenn Research Center, ONERA, and the University of Illinois are conducting a major research program whose goal is to improve our understanding of the aerodynamic scaling of ice accretions on airfoils. The program when it is completed will result in validated scaled simulation methods that produce the essential aerodynamic features of the full-scale iced-airfoil. This research will provide some of the first, high-fidelity, full-scale, iced-airfoil aerodynamic data. An initial study classified ice accretions based on their aerodynamics into four types: roughness, streamwise ice, horn ice, and spanwise-ridge ice. Subscale testing using a NACA 23012 airfoil was performed in the NASA IRT and University of Illinois wind tunnel to better understand the aerodynamics of these ice types and to test various levels of ice simulation fidelity. These studies are briefly reviewed here and have been presented in more detail in other papers. Based on these results, full-scale testing at the ONERA F1 tunnel using cast ice shapes obtained from molds taken in the IRT will provide full-scale iced airfoil data from full-scale ice accretions. Using these data as a baseline, the final step is to validate the simulation methods in scale in the Illinois wind tunnel. Computational ice accretion methods including LEWICE and ONICE have been used to guide the experiments and are briefly described and results shown. When full-scale and simulation aerodynamic results are available, these data will be used to further develop computational tools. Thus the purpose of the paper is to present an overview of the program and key results to date.
Recent Experiments at the Gottingen Aerodynamic Institute
NASA Technical Reports Server (NTRS)
Ackeret, J
1925-01-01
This report presents the results of various experiments carried out at the Gottingen Aerodynamic Institute. These include: experiments with Joukowski wing profiles; experiments on an airplane model with a built-in motor and functioning propeller; and the rotating cylinder (Magnus Effect).
NASA Technical Reports Server (NTRS)
Morgan, H. L., Jr.; Paulson, J. W., Jr.
1979-01-01
An investigation was conducted in the Langley V/STOL tunnel to determine the static longitudinal and lateral-directional aerodynamic characteristics of an advanced high-aspect-ratio supercritical-wing transport model equipped with a full-span leading-edge slat and part-span double-slotted trailing-edge flaps. This wide-body transport model was also equipped with spoiler and aileron control surfaces, flow-through nacelles, landing gear, movable horizontal tails, and interchangeable wing tips with aspect ratios of 10 and 12. The model was tested with leading-edge slat and trailing-edge flap combinations representative of cruise, climb, takeoff, and landing wing configurations. The tests were conducted at free-stream conditions corresponding to Reynolds numbers (based on mean geometric chord) of 0.97 to 1.63 x 10 to the 6th power and corresponding Mach numbers of 0.12 to 0.20, through an angle-of-attack range of -2 deg to 24 deg and a sideslip-angle range of -10 deg to 5 deg.
Investigation of Factors Affecting Aerodynamic Performance of Nebulized Nanoemulsion
Kamali, Hosein; Abbasi, Shayan; Amini, Mohammad Ali; Amani, Amir
2016-01-01
This work aimed to prepare a nanoemulsion preparation containing budesonide and assess its aerodynamic behavior in comparison with suspension of budesonide. In-vitro aerodynamic performance of the corresponding micellar solution (ie. nanoemulsion preparation without oil) was investigated too. Nanoemulsions of almond oil containing budesonide, as a hydrophobic model drug molecule, were prepared and optimized. Then, the effect of variation of surfactant/co-surfactant concentration on the aerodynamic properties of the nebulized aerosol was studied. The results indicated that the most physically stable formulation makes the smallest aerodynamic size. The concentration of co-surfactant was also shown to be critical in determination of aerodynamic size. Furthermore, the optimized sample, with 3% w/w almond oil, 20% w/w Tween 80+Span 80 and 2% w/w ethanol showed a smaller MMAD in comparison with the commercially available suspension and the micellar solution. PMID:28243265
NASA Astrophysics Data System (ADS)
Lin, Guofeng
Large-amplitude forced oscillation data for an F-18 configuration are analyzed with two modeling methods: Fourier functional analysis to form the indicial integrals, and a generalized dynamic aerodynamic model for stability and control analysis. The indicial integral is first applied to calculate the pitch damping parameter for comparison with the conventional forced oscillation test. It is shown that the reduced frequency affects the damping much more strongly than the test amplitude. Using the indicial integral models in a flight simulation code for an F-18 configuration, it is found that the configuration with unsteady aerodynamics becomes unstable in pitch if the pitch rate is high, in contrast to the quasi-steady configuration which depends mainly on the instantaneous angle of attack. In a pitch-up maneuver in the post-stall regime the configuration with unsteady aerodynamics can stay at a high pitch attitude and angle of attack without losing altitude for a much longer duration than the quasi-steady model. However, the speed will decrease faster because of higher drag. The newly developed generalized dynamic aerodynamic model is of the nonlinear algebraic form with the coefficients being determined from a set of large amplitude oscillatory experimental data by using least-square fitting. The resulting model coefficients are functions of the reduced frequency and amplitude. The new aerodynamic models have been verified with data in harmonic oscillation with a smaller amplitude and in constant pitch-rate motions. The new algebraic models are especially useful in stability and control analysis, and are used in bifurcation analysis and control studies for the same F-18 HARV configuration. The results show significant differences in the equilibrium surfaces and dynamic stability. It is also shown that control gains developed with the conventional quasi-steady aerodynamic data may not be adequate when the effect of unsteady aerodynamics is significant. A numerical
NASA Technical Reports Server (NTRS)
Potter, J. Leith
1992-01-01
Means for relatively simple and quick procedures are examined for estimating aerodynamic coefficients of lifting reentry vehicles. The methods developed allow aerospace designers not only to evaluate the aerodynamics of specific shapes but also to optimize shapes under given constraints. The analysis was also studied of the effect of thermomolecular flow on pressures measured by an orifice near the nose of a Space Shuttle Orbiter at altitudes above 75 km. It was shown that pressures corrected for thermomolecular flow effect are in good agreement with values predicted by independent theoretical methods. An incidental product was the insight gained about the free molecular thermal accommodation coefficient applicable under 'real' conditions of high speed flow in the Earth's atmosphere. The results are presented as abstracts of referenced papers. One reference paper is presented in its entirety.
NASA Astrophysics Data System (ADS)
Cain, T.; Owen, R.; Walton, C.
2005-02-01
The scramjet flight test Hyshot-2, flew on the 30 July 2002. The programme, led by the University of Queensland, had the primary objective of obtaining supersonic combustion data in flight for comparison with measurements made in shock tunnels. QinetiQ was one of the sponsors, and also provided aerodynamic data and trajectory predictions for the ballistic re-entry of the spinning sounding rocket. The unconventional missile geometry created by the nose-mounted asymmetric-scramjet in conjunction with the high angle of attack during re-entry makes the problem interesting. This paper presents the wind tunnel measurements and aerodynamic calculations used as input for the trajectory prediction. Indirect comparison is made with data obtained in the Hyshot-2 flight using a 6 degree-of-freedom trajectory simulation.
Advanced Aerodynamic Control Effectors
NASA Technical Reports Server (NTRS)
Wood, Richard M.; Bauer, Steven X. S.
1999-01-01
A 1990 research program that focused on the development of advanced aerodynamic control effectors (AACE) for military aircraft has been reviewed and summarized. Data are presented for advanced planform, flow control, and surface contouring technologies. The data show significant increases in lift, reductions in drag, and increased control power, compared to typical aerodynamic designs. The results presented also highlighted the importance of planform selection in the design of a control effector suite. Planform data showed that dramatic increases in lift (greater than 25%) can be achieved with multiple wings and a sawtooth forebody. Passive porosity and micro drag generator control effector data showed control power levels exceeding that available from typical effectors (moving surfaces). Application of an advanced planform to a tailless concept showed benefits of similar magnitude as those observed in the generic studies.
Aerodynamic Leidenfrost effect
NASA Astrophysics Data System (ADS)
Gauthier, Anaïs; Bird, James C.; Clanet, Christophe; Quéré, David
2016-12-01
When deposited on a plate moving quickly enough, any liquid can levitate as it does when it is volatile on a very hot solid (Leidenfrost effect). In the aerodynamic Leidenfrost situation, air gets inserted between the liquid and the moving solid, a situation that we analyze. We observe two types of entrainment. (i) The thickness of the air gap is found to increase with the plate speed, which is interpreted in the Landau-Levich-Derjaguin frame: Air is dynamically dragged along the surface and its thickness results from a balance between capillary and viscous effects. (ii) Air set in motion by the plate exerts a force on the levitating liquid. We discuss the magnitude of this aerodynamic force and show that it can be exploited to control the liquid and even to drive it against gravity.
Orion Crew Module Aerodynamic Testing
NASA Technical Reports Server (NTRS)
Murphy, Kelly J.; Bibb, Karen L.; Brauckmann, Gregory J.; Rhode, Matthew N.; Owens, Bruce; Chan, David T.; Walker, Eric L.; Bell, James H.; Wilson, Thomas M.
2011-01-01
The Apollo-derived Orion Crew Exploration Vehicle (CEV), part of NASA s now-cancelled Constellation Program, has become the reference design for the new Multi-Purpose Crew Vehicle (MPCV). The MPCV will serve as the exploration vehicle for all near-term human space missions. A strategic wind-tunnel test program has been executed at numerous facilities throughout the country to support several phases of aerodynamic database development for the Orion spacecraft. This paper presents a summary of the experimental static aerodynamic data collected to-date for the Orion Crew Module (CM) capsule. The test program described herein involved personnel and resources from NASA Langley Research Center, NASA Ames Research Center, NASA Johnson Space Flight Center, Arnold Engineering and Development Center, Lockheed Martin Space Sciences, and Orbital Sciences. Data has been compiled from eight different wind tunnel tests in the CEV Aerosciences Program. Comparisons are made as appropriate to highlight effects of angle of attack, Mach number, Reynolds number, and model support system effects.
Aerodynamics of the hovering hummingbird.
Warrick, Douglas R; Tobalske, Bret W; Powers, Donald R
2005-06-23
Despite profound musculoskeletal differences, hummingbirds (Trochilidae) are widely thought to employ aerodynamic mechanisms similar to those used by insects. The kinematic symmetry of the hummingbird upstroke and downstroke has led to the assumption that these halves of the wingbeat cycle contribute equally to weight support during hovering, as exhibited by insects of similar size. This assumption has been applied, either explicitly or implicitly, in widely used aerodynamic models and in a variety of empirical tests. Here we provide measurements of the wake of hovering rufous hummingbirds (Selasphorus rufus) obtained with digital particle image velocimetry that show force asymmetry: hummingbirds produce 75% of their weight support during the downstroke and only 25% during the upstroke. Some of this asymmetry is probably due to inversion of their cambered wings during upstroke. The wake of hummingbird wings also reveals evidence of leading-edge vortices created during the downstroke, indicating that they may operate at Reynolds numbers sufficiently low to exploit a key mechanism typical of insect hovering. Hummingbird hovering approaches that of insects, yet remains distinct because of effects resulting from an inherently dissimilar-avian-body plan.
Perching aerodynamics and trajectory optimization
NASA Astrophysics Data System (ADS)
Wickenheiser, Adam; Garcia, Ephrahim
2007-04-01
Advances in smart materials, actuators, and control architecture have enabled new flight capabilities for aircraft. Perching is one such capability, described as a vertical landing maneuver using in-flight shape reconfiguration in lieu of high thrust generation. A morphing, perching aircraft design is presented that is capable of post stall flight and very slow landing on a vertical platform. A comprehensive model of the aircraft's aerodynamics, with special regard to nonlinear affects such as flow separation and dynamic stall, is discussed. Trajectory optimization using nonlinear programming techniques is employed to show the effects that morphing and nonlinear aerodynamics have on the maneuver. These effects are shown to decrease the initial height and distance required to initiate the maneuver, reduce the bounds on the trajectory, and decrease the required thrust for the maneuver. Perching trajectories comparing morphing versus fixed-configuration and stalled versus un-stalled aircraft are presented. It is demonstrated that a vertical landing is possible in the absence of high thrust if post-stall flight capabilities and vehicle reconfiguration are utilized.
NASA Technical Reports Server (NTRS)
Cole, Jennifer Hansen
2010-01-01
This slide presentation reviews some of the basic principles of aerodynamics. Included in the presentation are: a few demonstrations of the principles, an explanation of the concepts of lift, drag, thrust and weight, a description of Bernoulli's principle, the concept of the airfoil (i.e., the shape of the wing) and how that effects lift, and the method of controlling an aircraft by manipulating the four forces using control surfaces.
NASA Technical Reports Server (NTRS)
Kelly, Thomas C.
1961-01-01
Aerodynamic loads results have been obtained in the Langley 8-foot transonic pressure tunnel at Mach numbers from 0.80 to 1.20 for a 1/10-scale model of the upper three stages of the Scout vehicle. Tests were conducted through an angle-of-attack range from -8 deg to 8 deg at an average test Reynolds number per foot of about 4.0 x 10(exp 6). Results indicated that the peak negative pressures associated with expansion corners at the nose and transition flare exhibit sizeable variations which occur over a relatively small Mach number range. The magnitude of the variations may cause the critical local loading condition for the full-scale vehicle to occur at a Mach number considerably lower than that at which the maximum dynamic pressure occurs in flight. The addition of protuberances simulating antennas and wiring conduits had slight, localized effects. The lift carryover from the nose and transition flare on the cylindrical portions of the model generally increased with an increase in Mach number.
Aerodynamic control in compressible flow using microwave driven discharges
NASA Astrophysics Data System (ADS)
McAndrew, Brendan
A new aerodynamic control scheme based on heating of the free stream flow is developed. The design, construction, and operation of a unique small scale wind tunnel to perform experiments involving this control scheme is detailed. Free stream heating is achieved by means of microwave driven discharges, and the resulting flow perturbations are used to alter the pressure distribution around a model in the flow. The experimental facility is also designed to allow the injection of an electron beam into the free stream for control of the discharge. Appropriate models for the fluid flow and discharge physics are developed, and comparisons of calculations based on those models are made with experimental results. The calculations have also been used to explore trends in parameters beyond the range possible in the experiments. The results of this work have been (1) the development of an operating facility capable of supporting free stream heat addition experiments in supersonic flow, (2) the development of a compatible instrumented model designed to make lift and drag measurements in a low pressure, high electrical noise environment, (3) a theoretical model to predict the change in breakdown threshold in the presence of an electron beam or other source of ionization, and (4) successful demonstration of aerodynamic control using free stream heat addition.
Aerodynamics and thermal physics of helicopter ice accretion
NASA Astrophysics Data System (ADS)
Han, Yiqiang
developed based on a set of 82 experimental measurements and also compared to existing predictions tools. Two reference predictions found in the literature yielded 76% and 54% discrepancy with respect to experimental testing, whereas the proposed ice roughness prediction model resulted in a 31% minimum accuracy in prediction. It must be noted that the accuracy of the proposed model is within the ice shape reproduction uncertainty of icing facilities. Based on the new ice roughness prediction model and the CSR heat transfer scaling method, an icing heat transfer model was developed. The approach achieved high accuracy in heat transfer prediction compared to experiments conducted at the AERTS facility. The discrepancy between predictions and experimental results was within +/-15%, which was within the measurement uncertainty range of the facility. By combining both the ice roughness and heat transfer predictions, and incorporating the modules into an existing ice prediction tool (LEWICE), improved prediction capability was obtained, especially for the glaze regime. With the available ice shapes accreted at the AERTS facility and additional experiments found in the literature, 490 sets of experimental ice shapes and corresponding aerodynamics testing data were available. A physics-based performance degradation empirical tool was developed and achieved a mean absolute deviation of 33% when compared to the entire experimental dataset, whereas 60% to 243% discrepancies were observed using legacy drag penalty prediction tools. Rotor torque predictions coupling Blade Element Momentum Theory and the proposed drag performance degradation tool was conducted on a total of 17 validation cases. The coupled prediction tool achieved a 10% predicting error for clean rotor conditions, and 16% error for iced rotor conditions. It was shown that additional roughness element could affect the measured drag by up to 25% during experimental testing, emphasizing the need of realistic ice structures
Development of an engineering code for the implementation of aerodynamic control devices in BEM
NASA Astrophysics Data System (ADS)
Aparicio, M.; González, A.; Gomez-Iradi, S.; Munduate, X.
2016-09-01
Aeroelastic codes based on Blade Element Momentum theory are the standard used by many wind turbine designers. These codes usually include models and corrections for unsteady aerodynamics, tip and root effect, tower shadow and other effects. In general, this kind of codes does not include models to correctly simulate aerodynamic control devices. This paper presents some modifications including the unsteady contributions due to the flap motion (based on indicial models) and the spanwise (3D) effects (based on circulation theory), in order to simulate flaps in the blades. This method can be included in BEM codes in general and it could also be applied to another kind of control devices. The validation and verification show the accuracy of this method using experimental data for two-dimensional unsteady cases, and CFD for three-dimensional steady and unsteady cases.
Aerodynamic Interaction Effects of a Helicopter Rotor and Fuselage
NASA Technical Reports Server (NTRS)
Boyd, David D., Jr.
1999-01-01
A three year Cooperative Research Agreements made in each of the three years between the Subsonic Aerodynamics Branch of the NASA Langley Research Center and the Virginia Polytechnic Institute and State University (Va. Tech) has been completed. This document presents results from this three year endeavor. The goal of creating an efficient method to compute unsteady interactional effects between a helicopter rotor and fuselage has been accomplished. This paper also includes appendices to support these findings. The topics are: 1) Rotor-Fuselage Interactions Aerodynamics: An Unsteady Rotor Model; and 2) Rotor/Fuselage Unsteady Interactional Aerodynamics: A New Computational Model.
Dynamic control of aerodynamic forces on a moving platform using active flow control
NASA Astrophysics Data System (ADS)
Brzozowski, Daniel P.
The unsteady interaction between trailing edge aerodynamic flow control and airfoil motion in pitch and plunge is investigated in wind tunnel experiments using a two degree-of-freedom 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 trailing edge of the airfoil using hybrid synthetic jet actuators. The dynamic coupling between the actuation and the time-dependent flow field is characterized using simultaneous force and particle image velocimetry (PIV) 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 both trajectory tracking and disturbance rejection maneuvers. The time-varying aerodynamic lift and pitching moment are estimated from a PIV wake survey using a reduced order model based on classical unsteady aerodynamic theory. These measurements suggest that the entire flow over the airfoil readjusts within 2--3 convective time scales, 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.
NASA Technical Reports Server (NTRS)
Hoad, D. R.
1975-01-01
A wind-tunnel investigation was conducted to determine the effect of deflecting the engine exit of a four-engine double-slotted flap transport to provide STOL performance. Longitudinal aerodynamic data were obtained at various engine exit positions and deflections. The data were obtained at three flap deflections representing cruise, take-off, and landing conditions for a range of angles of attack and various thrust coefficients. Downwash angles at the location of the horizontal tail were measured. The data are presented without analysis or discussion. Photographs of the test configurations are shown.
Aerodynamic Ground Effect in Fruitfly Sized Insect Takeoff
Kolomenskiy, Dmitry; Maeda, Masateru; Engels, Thomas; Liu, Hao; Schneider, Kai; Nave, Jean-Christophe
2016-01-01
Aerodynamic ground effect in flapping-wing insect flight is of importance to comparative morphologies and of interest to the micro-air-vehicle (MAV) community. Recent studies, however, show apparently contradictory results of either some significant extra lift or power savings, or zero ground effect. Here we present a numerical study of fruitfly sized insect takeoff with a specific focus on the significance of leg thrust and wing kinematics. Flapping-wing takeoff is studied using numerical modelling and high performance computing. The aerodynamic forces are calculated using a three-dimensional Navier–Stokes solver based on a pseudo-spectral method with volume penalization. It is coupled with a flight dynamics solver that accounts for the body weight, inertia and the leg thrust, while only having two degrees of freedom: the vertical and the longitudinal horizontal displacement. The natural voluntary takeoff of a fruitfly is considered as reference. The parameters of the model are then varied to explore possible effects of interaction between the flapping-wing model and the ground plane. These modified takeoffs include cases with decreased leg thrust parameter, and/or with periodic wing kinematics, constant body pitch angle. The results show that the ground effect during natural voluntary takeoff is negligible. In the modified takeoffs, when the rate of climb is slow, the difference in the aerodynamic forces due to the interaction with the ground is up to 6%. Surprisingly, depending on the kinematics, the difference is either positive or negative, in contrast to the intuition based on the helicopter theory, which suggests positive excess lift. This effect is attributed to unsteady wing-wake interactions. A similar effect is found during hovering. PMID:27019208
Low Speed Aerodynamics of the X-38 CRV
NASA Technical Reports Server (NTRS)
Komerath, N. M.; Funk, R.; Ames, R. G.; Mahalingam, R.; Matos, C.
1998-01-01
This project was performed in support of the engineering development of the NASA X-38 Crew Return Vehicle (CRV)system. Wind tunnel experiments were used to visualize various aerodynamic phenomena encountered by the CRV during the final stages of descent and landing. Scale models of the CRV were used to visualize vortex structures above and below the vehicle, and in its wake, and to quantify their trajectories. The effect of flaperon deflection on these structures was studied. The structure and dynamics of the CRV's wake during the drag parachute deployment stage were measured. Regions of high vorticity were identified using surveys conducted in several planes using a vortex meter. Periodic shedding of the vortex sheets from the sides of the CRV was observed using laser sheet videography as the CRV reached high angles of attack during the quasi-steady pitch-up prior to parafoil deployment. Using spectral analysis of hot-film anemometer data, the Strouhal number of these wake fluctuations was found to be 0.14 based on the model span. Phenomena encountered in flight test during parafoil operation were captured in scale-model tests, and a video photogrammetry technique was implemented to obtain parafoil surface shapes during flight in the tunnel. Forces on the parafoil were resolved using tension gages on individual lines. The temporal evolution of the phenomenon of leading edge collapse was captured. Laser velocimetry was used to demonstrate measurement of the porosity of the parafoil surface. From these measurements, several physical explanations have been developed for phenomena observed at various stages of the X-38 development program. Quantitative measurement capabilities have also been demonstrated for continued refinement of the aerodynamic technologies employed in the X-38 project. Detailed results from these studies are given in an AIAA Paper, two slide presentations, and other material which are given on a Web-based archival resource. This is the Digital
NASA Technical Reports Server (NTRS)
Sinha, Neeraj; Brinckman, Kevin; Jansen, Bernard; Seiner, John
2011-01-01
A method was developed of obtaining propulsive base flow data in both hot and cold jet environments, at Mach numbers and altitude of relevance to NASA launcher designs. The base flow data was used to perform computational fluid dynamics (CFD) turbulence model assessments of base flow predictive capabilities in order to provide increased confidence in base thermal and pressure load predictions obtained from computational modeling efforts. Predictive CFD analyses were used in the design of the experiments, available propulsive models were used to reduce program costs and increase success, and a wind tunnel facility was used. The data obtained allowed assessment of CFD/turbulence models in a complex flow environment, working within a building-block procedure to validation, where cold, non-reacting test data was first used for validation, followed by more complex reacting base flow validation.
NASA Technical Reports Server (NTRS)
Gillins, R. L.
1975-01-01
Force and moment data are presented which were obtained for each vehicle separately at a Mach number of 0.6, and for the mated orbiter/747 configuration at Mach numbers of 0.3, 0.5, 0.6, and 0.7. Orbiter angles of attack from 0 degrees to +12 degrees and 747/Carrier angles of attack from -3 degrees to +7 degrees were investigated at angles of sideslip of 0 degrees and -5 degrees. Model variables include orbiter elevon and rudder deflections, orbiter tail cone-on and off, various orbiter/747 attach structure configurations, 747 stabilizer and rudder deflections, and 747 CAM modification components-on and off. Photographs of test configurations are included.
NASA Technical Reports Server (NTRS)
Monta, W. J.
1980-01-01
The effects of conventional and square stores on the longitudinal aerodynamic characteristics of a fighter aircraft configuration at Mach numbers of 1.6, 1.8, and 2.0 was investigated. Five conventional store configurations and six arrangements of a square store configuration were studied. All configurations of the stores produced small, positive increments in the pitching moment throughout the angle-of-attack range, but the configuration with area ruled wing tanks also had a slight decrease on stability at the higher angles of attack. There were some small changes in lift coefficient because of the addition of the stores, causing the drag increment to vary with the lift coefficient. As a result, there were corresponding changes in the increments of the maximum lift drag ratios. The store drag coefficient based on the cross sectional area of the stores ranged from a maximum of 1.1 for the configuration with three Maverick missiles to a minimum of about .040 for the two MK-84 bombs and the arrangements with four square stores touching or two square stores in tandem. Square stores located side by side yielded about 0.50 in the aft position compared to 0.74 in the forward position.
ERIC Educational Resources Information Center
Hirsh, Alon; Levy, Sharona T.
2013-01-01
The present research addresses a curious finding: how learning physical principles enhanced athletes' biking performance but not their conceptual understanding. The study involves a model-based triathlon training program, Biking with Particles, concerning aerodynamics of biking in groups (drafting). A conceptual framework highlights several…
Aerodynamic characteristics of reentry vehicles at supersonic velocities
NASA Astrophysics Data System (ADS)
Adamov, N. P.; Kharitonov, A. M.; Chasovnikov, E. A.; Dyad'kin, A. A.; Kazakov, M. I.; Krylov, A. N.; Skorovarov, A. Yu.
2015-09-01
Models of promising reentry vehicles, experimental equipment, and test program are described. The method used to determine the total aerodynamic characteristics of these models on the AB-313 mechanical balance in the T-313 supersonic wind tunnel and the method used for simulations are presented. The aerodynamic coefficients of the examined objects in wide ranges of Mach numbers and angles of attack are obtained. The experimental data are compared with the results of simulations.
NASA Astrophysics Data System (ADS)
van der Male, Pim; van Dalen, Karel N.; Metrikine, Andrei V.
2016-11-01
Existing models for the analysis of offshore wind turbines account for the aerodynamic action on the turbine rotor in detail, requiring a high computational price. When considering the foundation of an offshore wind turbine, however, a reduced rotor model may be sufficient. To define such a model, the significance of the nonlinear velocity and history dependency of the aerodynamic force on a rotating blade should be known. Aerodynamic interaction renders the dynamics of a rotating blade in an ambient wind field nonlinear in terms of the dependency on the wind velocity relative to the structural motion. Moreover, the development in time of the aerodynamic force does not follow the flow velocity instantaneously, implying a history dependency. In addition, both the non-uniform blade geometry and the aerodynamic interaction couple the blade motions in and out of the rotational plane. Therefore, this study presents the Euler-Bernoulli formulation of a twisted rotating blade connected to a rigid hub, excited by either instantaneous or history-dependent aerodynamic forces. On this basis, the importance of the history dependency is determined. Moreover, to assess the nonlinear contributions, both models are linearized. The structural response is computed for a stand-still and a rotating blade, based on the NREL 5-MW turbine. To this end, the model is reduced on the basis of its first three free-vibration mode shapes. Blade tip response predictions, computed from turbulent excitation, correctly account for both modal and directional couplings, and the added damping resulting from the dependency of the aerodynamic force on the structural motion. Considering the deflection of the blade tip, the history-dependent and the instantaneous force models perform equally well, providing a basis for the potential use of the instantaneous model for the rotor reduction. The linearized instantaneous model provides similar results for the rotating blade, indicating its potential
NASA Astrophysics Data System (ADS)
Yongfeng, DENG; Jian, JIANG; Xianwei, HAN; Chang, TAN; Jianguo, WEI
2017-04-01
The problem of flow active control by low temperature plasma is considered to be one of the most flourishing fields of aerodynamics due to its practical advantages. Compared with other means, the electron beam plasma is a potential flow control method for large scale flow. In this paper, a computational fluid dynamics model coupled with a multi-fluid plasma model is established to investigate the aerodynamic characteristics induced by electron beam plasma. The results demonstrate that the electron beam strongly influences the flow properties, not only in the boundary layers, but also in the main flow. A weak shockwave is induced at the electron beam injection position and develops to the other side of the wind tunnel behind the beam. It brings additional energy into air, and the inducing characteristics are closely related to the beam power and increase nonlinearly with it. The injection angles also influence the flow properties to some extent. Based on this research, we demonstrate that the high energy electron beam air plasma has three attractive advantages in aerodynamic applications, i.e. the high energy density, wide action range and excellent action effect. Due to the rapid development of near space hypersonic vehicles and atmospheric fighters, by optimizing the parameters, the electron beam can be used as an alternative means in aerodynamic steering in these applications.
Nozzle Aerodynamic Stability During a Throat Shift
NASA Technical Reports Server (NTRS)
Kawecki, Edwin J.; Ribeiro, Gregg L.
2005-01-01
An experimental investigation was conducted on the internal aerodynamic stability of a family of two-dimensional (2-D) High Speed Civil Transport (HSCT) nozzle concepts. These nozzles function during takeoff as mixer-ejectors to meet acoustic requirements, and then convert to conventional high-performance convergent-divergent (CD) nozzles at cruise. The transition between takeoff mode and cruise mode results in the aerodynamic throat and the minimum cross-sectional area that controls the engine backpressure shifting location within the nozzle. The stability and steadiness of the nozzle aerodynamics during this so called throat shift process can directly affect the engine aerodynamic stability, and the mechanical design of the nozzle. The objective of the study was to determine if pressure spikes or other perturbations occurred during the throat shift process and, if so, identify the caused mechanisms for the perturbations. The two nozzle concepts modeled in the test program were the fixed chute (FC) and downstream mixer (DSM). These 2-D nozzles differ principally in that the FC has a large over-area between the forward throat and aft throat locations, while the DSM has an over-area of only about 10 percent. The conclusions were that engine mass flow and backpressure can be held constant simultaneously during nozzle throat shifts on this class of nozzles, and mode shifts can be accomplished at a constant mass flow and engine backpressure without upstream pressure perturbations.
Freight Wing Trailer Aerodynamics
Graham, Sean; Bigatel, Patrick
2004-10-17
Freight Wing Incorporated utilized the opportunity presented by this DOE category one Inventions and Innovations grant to successfully research, develop, test, patent, market, and sell innovative fuel and emissions saving aerodynamic attachments for the trucking industry. A great deal of past scientific research has demonstrated that streamlining box shaped semi-trailers can significantly reduce a truck's fuel consumption. However, significant design challenges have prevented past concepts from meeting industry needs. Market research early in this project revealed the demands of truck fleet operators regarding aerodynamic attachments. Products must not only save fuel, but cannot interfere with the operation of the truck, require significant maintenance, add significant weight, and must be extremely durable. Furthermore, SAE/TMC J1321 tests performed by a respected independent laboratory are necessary for large fleets to even consider purchase. Freight Wing used this information to create a system of three practical aerodynamic attachments for the front, rear and undercarriage of standard semi trailers. SAE/TMC J1321 Type II tests preformed by the Transportation Research Center (TRC) demonstrated a 7% improvement to fuel economy with all three products. If Freight Wing is successful in its continued efforts to gain market penetration, the energy and environmental savings would be considerable. Each truck outfitted saves approximately 1,100 gallons of fuel every 100,000 miles, which prevents over 12 tons of CO2 from entering the atmosphere. If all applicable trailers used the technology, the country could save approximately 1.8 billion gallons of diesel fuel, 18 million tons of emissions and 3.6 billion dollars annually.
TAD- THEORETICAL AERODYNAMICS PROGRAM
NASA Technical Reports Server (NTRS)
Barrowman, J.
1994-01-01
This theoretical aerodynamics program, TAD, was developed to predict the aerodynamic characteristics of vehicles with sounding rocket configurations. These slender, axisymmetric finned vehicle configurations have a wide range of aeronautical applications from rockets to high speed armament. Over a given range of Mach numbers, TAD will compute the normal force coefficient derivative, the center-of-pressure, the roll forcing moment coefficient derivative, the roll damping moment coefficient derivative, and the pitch damping moment coefficient derivative of a sounding rocket configured vehicle. The vehicle may consist of a sharp pointed nose of cone or tangent ogive shape, up to nine other body divisions of conical shoulder, conical boattail, or circular cylinder shape, and fins of trapezoid planform shape with constant cross section and either three or four fins per fin set. The characteristics computed by TAD have been shown to be accurate to within ten percent of experimental data in the supersonic region. The TAD program calculates the characteristics of separate portions of the vehicle, calculates the interference between separate portions of the vehicle, and then combines the results to form a total vehicle solution. Also, TAD can be used to calculate the characteristics of the body or fins separately as an aid in the design process. Input to the TAD program consists of simple descriptions of the body and fin geometries and the Mach range of interest. Output includes the aerodynamic characteristics of the total vehicle, or user-selected portions, at specified points over the mach range. The TAD program is written in FORTRAN IV for batch execution and has been implemented on an IBM 360 computer with a central memory requirement of approximately 123K of 8 bit bytes. The TAD program was originally developed in 1967 and last updated in 1972.