Review of design optimization methods for turbomachinery aerodynamics

NASA Astrophysics Data System (ADS)

Li, Zhihui; Zheng, Xinqian

2017-08-01

In today's competitive environment, new turbomachinery designs need to be not only more efficient, quieter, and ;greener; but also need to be developed at on much shorter time scales and at lower costs. A number of advanced optimization strategies have been developed to achieve these requirements. This paper reviews recent progress in turbomachinery design optimization to solve real-world aerodynamic problems, especially for compressors and turbines. This review covers the following topics that are important for optimizing turbomachinery designs. (1) optimization methods, (2) stochastic optimization combined with blade parameterization methods and the design of experiment methods, (3) gradient-based optimization methods for compressors and turbines and (4) data mining techniques for Pareto Fronts. We also present our own insights regarding the current research trends and the future optimization of turbomachinery designs.

Venusian atmospheric and Magellan properties from attitude control data. M.S. Thesis

NASA Technical Reports Server (NTRS)

Croom, Christopher A.; Tolson, Robert H.

1994-01-01

Results are presented of the study of the Venusian atmosphere, Magellan aerodynamic moment coefficients, moments of inertia, and solar moment coefficients. This investigation is based upon the use of attitude control data in the form of reaction wheel speeds from the Magellan spacecraft. As the spacecraft enters the upper atmosphere of Venus, measurable torques are experienced due to aerodynamic effects. Solar and gravity gradient effects also cause additional torques throughout the orbit. In order to maintain an inertially fixed attitude, the control system counteracts these torques by changing the angular rates of three reaction wheels. Model reaction wheel speeds are compared to observed Magellan reaction wheel speeds through a differential correction procedure. This method determines aerodynamic, atmospheric, solar pressure, and mass moment of inertia parameters. Atmospheric measurements include both base densities and scale heights. Atmospheric base density results confirm natural variability as measured by the standard orbital decay method. Potential inconsistencies in free molecular aerodynamic moment coefficients are identified. Moments of inertia are determined with a precision better than 1 percent of the largest principal moment of inertia.

Reentry-Vehicle Shape Optimization Using a Cartesian Adjoint Method and CAD Geometry

NASA Technical Reports Server (NTRS)

Nemec, Marian; Aftosmis, Michael J.

2006-01-01

A DJOINT solutions of the governing flow equations are becoming increasingly important for the development of efficient analysis and optimization algorithms. A well-known use of the adjoint method is gradient-based shape. Given an objective function that defines some measure of performance, such as the lift and drag functionals, its gradient is computed at a cost that is essentially independent of the number of design variables (e.g., geometric parameters that control the shape). Classic aerodynamic applications of gradient-based optimization include the design of cruise configurations for transonic and supersonic flow, as well as the design of high-lift systems. are perhaps the most promising approach for addressing the issues of flow solution automation for aerodynamic design problems. In these methods, the discretization of the wetted surface is decoupled from that of the volume mesh. This not only enables fast and robust mesh generation for geometry of arbitrary complexity, but also facilitates access to geometry modeling and manipulation using parametric computer-aided design (CAD). In previous work on Cartesian adjoint solvers, Melvin et al. developed an adjoint formulation for the TRANAIR code, which is based on the full-potential equation with viscous corrections. More recently, Dadone and Grossman presented an adjoint formulation for the two-dimensional Euler equations using a ghost-cell method to enforce the wall boundary conditions. In Refs. 18 and 19, we presented an accurate and efficient algorithm for the solution of the adjoint Euler equations discretized on Cartesian meshes with embedded, cut-cell boundaries. Novel aspects of the algorithm were the computation of surface shape sensitivities for triangulations based on parametric-CAD models and the linearization of the coupling between the surface triangulation and the cut-cells. The accuracy of the gradient computation was verified using several three-dimensional test cases, which included design variables such as the free stream parameters and the planform shape of an isolated wing. The objective of the present work is to extend our adjoint formulation to problems involving general shape changes. Factors under consideration include the computation of mesh sensitivities that provide a reliable approximation of the objective function gradient, as well as the computation of surface shape sensitivities based on a direct-CAD interface. We present detailed gradient verification studies and then focus on a shape optimization problem for an Apollo-like reentry vehicle. The goal of the optimization is to enhance the lift-to-drag ratio of the capsule by modifying the shape of its heat-shield in conjunction with a center-of-gravity (c.g.) offset. This multipoint and multi-objective optimization problem is used to demonstrate the overall effectiveness of the Cartesian adjoint method for addressing the issues of complex aerodynamic design.

Modeling coupled aerodynamics and vocal fold dynamics using immersed boundary methods.

PubMed

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.

Thermal lift generation and drag reduction in rarefied aerodynamics

NASA Astrophysics Data System (ADS)

Pekardan, Cem; Alexeenko, Alina

2016-11-01

With the advent of the new technologies in low pressure environments such as Hyperloop and helicopters designed for Martian applications, understanding the aerodynamic behavior of airfoils in rarefied environments are becoming more crucial. In this paper, verification of rarefied ES-BGK solver and ideas such as prediction of the thermally induced lift and drag reduction in rarefied aerodynamics are investigated. Validation of the rarefied ES-BGK solver with Runge-Kutta discontinous Galerkin method with experiments in transonic regime with a Reynolds number of 73 showed that ES-BGK solver is the most suitable solver in near slip transonic regime. For the quantification of lift generation, A NACA 0012 airfoil is studied with a high temperature surface on the bottom for the lift creation for different Knudsen numbers. It was seen that for lower velocities, continuum solver under predicts the lift generation when the Knudsen number is 0.00129 due to local velocity gradients reaching slip regime although lift coefficient is higher with the Boltzmann ES-BGK solutions. In the second part, the feasibility of using thermal transpiration for drag reduction is studied. Initial study in drag reduction includes an application of a thermal gradient at the upper surface of a NACA 0012 airfoil near trailing edge at a 12-degree angle of attack and 5 Pa pressure. It was seen that drag is reduced by 4 percent and vortex shedding frequency is reduced due to asymmetry introduced in the flow due to temperature gradient causing reverse flow due to thermal transpiration phenomena.

Three-Dimensional Viscous Alternating Direction Implicit Algorithm and Strategies for Shape Optimization

NASA Technical Reports Server (NTRS)

Pandya, Mohagna J.; Baysal, Oktay

1997-01-01

A gradient-based shape optimization based on quasi-analytical sensitivities has been extended for practical three-dimensional aerodynamic applications. The flow analysis has been rendered by a fully implicit, finite-volume formulation of the Euler and Thin-Layer Navier-Stokes (TLNS) equations. Initially, the viscous laminar flow analysis for a wing has been compared with an independent computational fluid dynamics (CFD) code which has been extensively validated. The new procedure has been demonstrated in the design of a cranked arrow wing at Mach 2.4 with coarse- and fine-grid based computations performed with Euler and TLNS equations. The influence of the initial constraints on the geometry and aerodynamics of the optimized shape has been explored. Various final shapes generated for an identical initial problem formulation but with different optimization path options (coarse or fine grid, Euler or TLNS), have been aerodynamically evaluated via a common fine-grid TLNS-based analysis. The initial constraint conditions show significant bearing on the optimization results. Also, the results demonstrate that to produce an aerodynamically efficient design, it is imperative to include the viscous physics in the optimization procedure with the proper resolution. Based upon the present results, to better utilize the scarce computational resources, it is recommended that, a number of viscous coarse grid cases using either a preconditioned bi-conjugate gradient (PbCG) or an alternating-direction-implicit (ADI) method, should initially be employed to improve the optimization problem definition, the design space and initial shape. Optimized shapes should subsequently be analyzed using a high fidelity (viscous with fine-grid resolution) flow analysis to evaluate their true performance potential. Finally, a viscous fine-grid-based shape optimization should be conducted, using an ADI method, to accurately obtain the final optimized shape.

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.

Aerodynamic Design Optimization on Unstructured Meshes Using the Navier-Stokes Equations

NASA Technical Reports Server (NTRS)

Nielsen, Eric J.; Anderson, W. Kyle

1998-01-01

A discrete adjoint method is developed and demonstrated for aerodynamic design optimization on unstructured grids. The governing equations are the three-dimensional Reynolds-averaged Navier-Stokes equations coupled with a one-equation turbulence model. A discussion of the numerical implementation of the flow and adjoint equations is presented. Both compressible and incompressible solvers are differentiated and the accuracy of the sensitivity derivatives is verified by comparing with gradients obtained using finite differences. Several simplifying approximations to the complete linearization of the residual are also presented, and the resulting accuracy of the derivatives is examined. Demonstration optimizations for both compressible and incompressible flows are given.

Aerodynamic optimization of supersonic compressor cascade using differential evolution on GPU

NASA Astrophysics Data System (ADS)

Aissa, Mohamed Hasanine; Verstraete, Tom; Vuik, Cornelis

2016-06-01

Differential Evolution (DE) is a powerful stochastic optimization method. Compared to gradient-based algorithms, DE is able to avoid local minima but requires at the same time more function evaluations. In turbomachinery applications, function evaluations are performed with time-consuming CFD simulation, which results in a long, non affordable, design cycle. Modern High Performance Computing systems, especially Graphic Processing Units (GPUs), are able to alleviate this inconvenience by accelerating the design evaluation itself. In this work we present a validated CFD Solver running on GPUs, able to accelerate the design evaluation and thus the entire design process. An achieved speedup of 20x to 30x enabled the DE algorithm to run on a high-end computer instead of a costly large cluster. The GPU-enhanced DE was used to optimize the aerodynamics of a supersonic compressor cascade, achieving an aerodynamic loss minimization of 20%.

Aerodynamic optimization of supersonic compressor cascade using differential evolution on GPU

DOE Office of Scientific and Technical Information (OSTI.GOV)

Aissa, Mohamed Hasanine; Verstraete, Tom; Vuik, Cornelis

Differential Evolution (DE) is a powerful stochastic optimization method. Compared to gradient-based algorithms, DE is able to avoid local minima but requires at the same time more function evaluations. In turbomachinery applications, function evaluations are performed with time-consuming CFD simulation, which results in a long, non affordable, design cycle. Modern High Performance Computing systems, especially Graphic Processing Units (GPUs), are able to alleviate this inconvenience by accelerating the design evaluation itself. In this work we present a validated CFD Solver running on GPUs, able to accelerate the design evaluation and thus the entire design process. An achieved speedup of 20xmore » to 30x enabled the DE algorithm to run on a high-end computer instead of a costly large cluster. The GPU-enhanced DE was used to optimize the aerodynamics of a supersonic compressor cascade, achieving an aerodynamic loss minimization of 20%.« less

Demonstration of Automatically-Generated Adjoint Code for Use in Aerodynamic Shape Optimization

NASA Technical Reports Server (NTRS)

Green, Lawrence; Carle, Alan; Fagan, Mike

1999-01-01

Gradient-based optimization requires accurate derivatives of the objective function and constraints. These gradients may have previously been obtained by manual differentiation of analysis codes, symbolic manipulators, finite-difference approximations, or existing automatic differentiation (AD) tools such as ADIFOR (Automatic Differentiation in FORTRAN). Each of these methods has certain deficiencies, particularly when applied to complex, coupled analyses with many design variables. Recently, a new AD tool called ADJIFOR (Automatic Adjoint Generation in FORTRAN), based upon ADIFOR, was developed and demonstrated. Whereas ADIFOR implements forward-mode (direct) differentiation throughout an analysis program to obtain exact derivatives via the chain rule of calculus, ADJIFOR implements the reverse-mode counterpart of the chain rule to obtain exact adjoint form derivatives from FORTRAN code. Automatically-generated adjoint versions of the widely-used CFL3D computational fluid dynamics (CFD) code and an algebraic wing grid generation code were obtained with just a few hours processing time using the ADJIFOR tool. The codes were verified for accuracy and were shown to compute the exact gradient of the wing lift-to-drag ratio, with respect to any number of shape parameters, in about the time required for 7 to 20 function evaluations. The codes have now been executed on various computers with typical memory and disk space for problems with up to 129 x 65 x 33 grid points, and for hundreds to thousands of independent variables. These adjoint codes are now used in a gradient-based aerodynamic shape optimization problem for a swept, tapered wing. For each design iteration, the optimization package constructs an approximate, linear optimization problem, based upon the current objective function, constraints, and gradient values. The optimizer subroutines are called within a design loop employing the approximate linear problem until an optimum shape is found, the design loop limit is reached, or no further design improvement is possible due to active design variable bounds and/or constraints. The resulting shape parameters are then used by the grid generation code to define a new wing surface and computational grid. The lift-to-drag ratio and its gradient are computed for the new design by the automatically-generated adjoint codes. Several optimization iterations may be required to find an optimum wing shape. Results from two sample cases will be discussed. The reader should note that this work primarily represents a demonstration of use of automatically- generated adjoint code within an aerodynamic shape optimization. As such, little significance is placed upon the actual optimization results, relative to the method for obtaining the results.

Aerodynamic Shape Sensitivity Analysis and Design Optimization of Complex Configurations Using Unstructured Grids

NASA Technical Reports Server (NTRS)

Taylor, Arthur C., III; Newman, James C., III; Barnwell, Richard W.

1997-01-01

A three-dimensional unstructured grid approach to aerodynamic shape sensitivity analysis and design optimization has been developed and is extended to model geometrically complex configurations. The advantage of unstructured grids (when compared with a structured-grid approach) is their inherent ability to discretize irregularly shaped domains with greater efficiency and less effort. Hence, this approach is ideally suited for geometrically complex configurations of practical interest. In this work the nonlinear Euler equations are solved using an upwind, cell-centered, finite-volume scheme. The discrete, linearized systems which result from this scheme are solved iteratively by a preconditioned conjugate-gradient-like algorithm known as GMRES for the two-dimensional geometry and a Gauss-Seidel algorithm for the three-dimensional; similar procedures are used to solve the accompanying linear aerodynamic sensitivity equations in incremental iterative form. As shown, this particular form of the sensitivity equation makes large-scale gradient-based aerodynamic optimization possible by taking advantage of memory efficient methods to construct exact Jacobian matrix-vector products. Simple parameterization techniques are utilized for demonstrative purposes. Once the surface has been deformed, the unstructured grid is adapted by considering the mesh as a system of interconnected springs. Grid sensitivities are obtained by differentiating the surface parameterization and the grid adaptation algorithms with ADIFOR (which is an advanced automatic-differentiation software tool). To demonstrate the ability of this procedure to analyze and design complex configurations of practical interest, the sensitivity analysis and shape optimization has been performed for a two-dimensional high-lift multielement airfoil and for a three-dimensional Boeing 747-200 aircraft.

A modified micrometeorological gradient method for estimating O3 dry depositions over a forest canopy

NASA Astrophysics Data System (ADS)

Wu, Z. Y.; Zhang, L.; Wang, X. M.; Munger, J. W.

2015-07-01

Small pollutant concentration gradients between levels above a plant canopy result in large uncertainties in estimated air-surface exchange fluxes when using existing micrometeorological gradient methods, including the aerodynamic gradient method (AGM) and the modified Bowen ratio method (MBR). A modified micrometeorological gradient method (MGM) is proposed in this study for estimating O3 dry deposition fluxes over a forest canopy using concentration gradients between a level above and a level below the canopy top, taking advantage of relatively large gradients between these levels due to significant pollutant uptake in the top layers of the canopy. The new method is compared with the AGM and MBR methods and is also evaluated using eddy-covariance (EC) flux measurements collected at the Harvard Forest Environmental Measurement Site, Massachusetts, during 1993-2000. All three gradient methods (AGM, MBR, and MGM) produced similar diurnal cycles of O3 dry deposition velocity (Vd(O3)) to the EC measurements, with the MGM method being the closest in magnitude to the EC measurements. The multi-year average Vd(O3) differed significantly between these methods, with the AGM, MBR, and MGM method being 2.28, 1.45, and 1.18 times that of the EC, respectively. Sensitivity experiments identified several input parameters for the MGM method as first-order parameters that affect the estimated Vd(O3). A 10% uncertainty in the wind speed attenuation coefficient or canopy displacement height can cause about 10% uncertainty in the estimated Vd(O3). An unrealistic leaf area density vertical profile can cause an uncertainty of a factor of 2.0 in the estimated Vd(O3). Other input parameters or formulas for stability functions only caused an uncertainly of a few percent. The new method provides an alternative approach to monitoring/estimating long-term deposition fluxes of similar pollutants over tall canopies.

Computer Software Management and Information Center

NASA Technical Reports Server (NTRS)

1983-01-01

Computer programs for passive anti-roll tank, earth resources laboratory applications, the NIMBUS-7 coastal zone color scanner derived products, transportable applications executive, plastic and failure analysis of composites, velocity gradient method for calculating velocities in an axisymmetric annular duct, an integrated procurement management system, data I/O PRON for the Motorola exorcisor, aerodynamic shock-layer shape, kinematic modeling, hardware library for a graphics computer, and a file archival system are documented.

UNAERO: A package of FORTRAN subroutines for approximating unsteady aerodynamics in the time domain

NASA Technical Reports Server (NTRS)

Dunn, H. J.

1985-01-01

This report serves as an instruction and maintenance manual for a collection of CDC CYBER FORTRAN IV subroutines for approximating the unsteady aerodynamic forces in the time domain. The result is a set of constant-coefficient first-order differential equations that approximate the dynamics of the vehicle. Provisions are included for adjusting the number of modes used for calculating the approximations so that an accurate approximation is generated. The number of data points at different values of reduced frequency can also be varied to adjust the accuracy of the approximation over the reduced-frequency range. The denominator coefficients of the approximation may be calculated by means of a gradient method or a least-squares approximation technique. Both the approximation methods use weights on the residual error. A new set of system equations, at a different dynamic pressure, can be generated without the approximations being recalculated.

Conference on Low Reynolds Number Airfoil Aerodynamics, Notre Dame, IN, June 16-18, 1985, Proceedings

NASA Technical Reports Server (NTRS)

Mueller, T. J. (Editor)

1985-01-01

Topics of interest in the design, flow modeling and visualization, and turbulence and flow separation effects for low Reynolds number (Re) airfoils are discussed. Design methods are presented for Re from 50,000-500,000, including a viscous-inviscid coupling method and by using a constrained pitching moment. The effects of pressure gradients, unsteady viscous aerodynamics and separation bubbles are investigated, with particular note made of factors which most influence the size and location of separation bubbles and control their effects. Attention is also given to experimentation with low Re airfoils and to numerical models of symmetry breaking and lift hysteresis from separation. Both steady and unsteady flow experiments are reviewed, with the trials having been held in wind tunnels and the free atmosphere. The topics discussed are of interest to designers of RPVs, high altitude aircraft, sailplanes, ultralights and wind turbines.

Aerodynamic Design of Complex Configurations Using Cartesian Methods and CAD Geometry

NASA Technical Reports Server (NTRS)

Nemec, Marian; Aftosmis, Michael J.; Pulliam, Thomas H.

2003-01-01

The objective for this paper is to present the development of an optimization capability for the Cartesian inviscid-flow analysis package of Aftosmis et al. We evaluate and characterize the following modules within the new optimization framework: (1) A component-based geometry parameterization approach using a CAD solid representation and the CAPRI interface. (2) The use of Cartesian methods in the development Optimization techniques using a genetic algorithm. The discussion and investigations focus on several real world problems of the optimization process. We examine the architectural issues associated with the deployment of a CAD-based design approach in a heterogeneous parallel computing environment that contains both CAD workstations and dedicated compute nodes. In addition, we study the influence of noise on the performance of optimization techniques, and the overall efficiency of the optimization process for aerodynamic design of complex three-dimensional configurations. of automated optimization tools. rithm and a gradient-based algorithm.

An Investigation of the Compatibility of Radiation and Convection Heat Flux Measurements

NASA Technical Reports Server (NTRS)

Liebert, Curt H.

1996-01-01

A method for determining time-resolved absorbed surface heat flux and surface temperature in radiation and convection environments is described. The method is useful for verification of aerodynamic, heat transfer and durability models. A practical heat flux gage fabrication procedure and a simple one-dimensional inverse heat conduction model and calculation procedure are incorporated in this method. The model provides an estimate of the temperature and heat flux gradient in the direction of heat transfer through the gage. This paper discusses several successful time-resolved tests of this method in hostile convective heating and cooling environments.

JPRS report: Science and technology. Central Eurasia: Engineering and equipment

NASA Astrophysics Data System (ADS)

1993-10-01

Translated articles cover the following topics: transient gas dynamic processes in ramjet engines; aerodynamic characteristics of delta wings with detached leading edge shock wave at hypersonic flight velocities; effect of atmospheric density gradient on aerodynamic stabilization; measurement of target radar scattering characteristics using frequency synthesized signals; assessing survivability and ensuring safety of large axial-flow compressor blades; procedure for experimentally determining transient aerodynamic forces caused by flat vane cascade; analysis of aerodynamic interaction of profile and vortex; laser machine for balancing dynamically adjusted gyros; use of heat pumps in solar heat supply systems; numerical simulation of deflagration transition to detonation in homogeneous combustible fuel mixture; and investigation of chemically nonequilibrium flow about bodies allowing for vibrational relaxation.

Adjoint Sensitivity Computations for an Embedded-Boundary Cartesian Mesh Method and CAD Geometry

NASA Technical Reports Server (NTRS)

Nemec, Marian; Aftosmis,Michael J.

2006-01-01

Cartesian-mesh methods are perhaps the most promising approach for addressing the issues of flow solution automation for aerodynamic design problems. In these methods, the discretization of the wetted surface is decoupled from that of the volume mesh. This not only enables fast and robust mesh generation for geometry of arbitrary complexity, but also facilitates access to geometry modeling and manipulation using parametric Computer-Aided Design (CAD) tools. Our goal is to combine the automation capabilities of Cartesian methods with an eficient computation of design sensitivities. We address this issue using the adjoint method, where the computational cost of the design sensitivities, or objective function gradients, is esseutially indepeudent of the number of design variables. In previous work, we presented an accurate and efficient algorithm for the solution of the adjoint Euler equations discretized on Cartesian meshes with embedded, cut-cell boundaries. Novel aspects of the algorithm included the computation of surface shape sensitivities for triangulations based on parametric-CAD models and the linearization of the coupling between the surface triangulation and the cut-cells. The objective of the present work is to extend our adjoint formulation to problems involving general shape changes. Central to this development is the computation of volume-mesh sensitivities to obtain a reliable approximation of the objective finction gradient. Motivated by the success of mesh-perturbation schemes commonly used in body-fitted unstructured formulations, we propose an approach based on a local linearization of a mesh-perturbation scheme similar to the spring analogy. This approach circumvents most of the difficulties that arise due to non-smooth changes in the cut-cell layer as the boundary shape evolves and provides a consistent approximation tot he exact gradient of the discretized abjective function. A detailed gradient accurace study is presented to verify our approach. Thereafter, we focus on a shape optimization problem for an Apollo-like reentry capsule. The optimization seeks to enhance the lift-to-drag ratio of the capsule by modifyjing the shape of its heat-shield in conjunction with a center-of-gravity (c.g.) offset. This multipoint and multi-objective optimization problem is used to demonstrate the overall effectiveness of the Cartesian adjoint method for addressing the issues of complex aerodynamic design. This abstract presents only a brief outline of the numerical method and results; full details will be given in the final paper.

A novel concept for subsonic inlet boundary-layer control

NASA Technical Reports Server (NTRS)

Miller, B. A.

1977-01-01

A self-bleeding method for boundary layer control is described and tested for a subsonic inlet designed to operate in the flowfield generated by high angles of attack. Naturally occurring surface static pressure gradients are used to remove the boundary layer from a separation-prone region of the inlet and to reinject it at a less critical location with a net performance gain. The results suggest that this self-bleeding method for boundary-layer control might be successfully applied to other inlets operating at extreme aerodynamic conditions.

High-Fidelity Aerodynamic Shape Optimization for Natural Laminar Flow

NASA Astrophysics Data System (ADS)

Rashad, Ramy

To ensure the long-term sustainability of aviation, serious effort is underway to mitigate the escalating economic, environmental, and social concerns of the industry. Significant improvement to the energy efficiency of air transportation is required through the research and development of advanced and unconventional airframe and engine technologies. In the quest to reduce airframe drag, this thesis is concerned with the development and demonstration of an effective design tool for improving the aerodynamic efficiency of subsonic and transonic airfoils. The objective is to advance the state-of-the-art in high-fidelity aerodynamic shape optimization by incorporating and exploiting the phenomenon of laminar-turbulent transition in an efficient manner. A framework for the design and optimization of Natural Laminar Flow (NLF) airfoils is developed and demonstrated with transition prediction capable of accounting for the effects of Reynolds number, freestream turbulence intensity, Mach number, and pressure gradients. First, a two-dimensional Reynolds-averaged Navier-Stokes (RANS) flow solver has been extended to incorporate an iterative laminar-turbulent transition prediction methodology. The natural transition locations due to Tollmien-Schlichting instabilities are predicted using the simplified eN envelope method of Drela and Giles or, alternatively, the compressible form of the Arnal-Habiballah-Delcourt criterion. The boundary-layer properties are obtained directly from the Navier-Stokes flow solution, and the transition to turbulent flow is modeled using an intermittency function in conjunction with the Spalart-Allmaras turbulence model. The RANS solver is subsequently employed in a gradient-based sequential quadratic programming shape optimization framework. The laminar-turbulent transition criteria are tightly coupled into the objective and gradient evaluations. The gradients are obtained using a new augmented discrete-adjoint formulation for non-local transition criteria. Using the eN transition criterion, the proposed framework is applied to the single and multipoint optimization of subsonic and transonic airfoils, leading to robust NLF designs. The aerodynamic design requirements over a range of cruise flight conditions are cast into a multipoint optimization problem through a composite objective defined using a weighted integral of the operating points. To study and quantify off-design performance, a Pareto front is formed using a weighted objective combining free-transition and fully-turbulent operating conditions. Next we examine the sensitivity of NLF design to the freestream disturbance environment, highlighting the on- and off-design performance at different critical N-factors. Finally, we propose and demonstrate a technique to enable the design of airfoils with robust performance over a range of critical N-factors.

Field comparison of an eddy accumulation and an aerodynamic-gradient system for measuring pesticide volatilization fluxes

USGS Publications Warehouse

Majewski, M.; Desjardina, R.; Rochette, P.; Pattey, E.; Selber, J.; Glotfelty, D.

1993-01-01

The field experiment reported here applied the relaxed eddy accumulation (REA) technique to the measurement of triallate (TA) and trifluralin (TF) volatilization from fallow soil. A critical analysis of the REA system used in this experiment is done, and the fluxes are compared to those obtained by the aerodynamic-gradient (AG) technique. The measured cumulative volatilization losses, corrected for the effective upwind source area (footprint), for the AG system were higher than with the REA system. The differences between the methods over the first 5 days of the experiment were 27 and 13% for TA and TF, respectively. A mass balance based on the amount of parent compounds volatilized from soil during the first 5 days of the experiment showed a 110 and 70% and a 79 and 61% accountability for triallate and trifluralin by the AG and REA methods, respectively. These results also show that the non-footprint-corrected AG flux values underestimated the volatilization flux by approximately 16%. The footprint correction model used in this experiment does not presently have the capability of accounting for changes in atmospheric stability. However, these values still provide an indication of the most likely upwind area affecting the evaporative flux estimations. The soil half-lives for triallate and trifluralin were 9.8 and 7.0 days, respectively. ?? 1992 American Chemical Society.

Improvement of the cruise performances of a wing by means of aerodynamic optimization. Validation with a Far-Field method

NASA Astrophysics Data System (ADS)

Jiménez-Varona, J.; Ponsin Roca, J.

2015-06-01

Under a contract with AIRBUS MILITARY (AI-M), an exercise to analyze the potential of optimization techniques to improve the wing performances at cruise conditions has been carried out by using an in-house design code. The original wing was provided by AI-M and several constraints were posed for the redesign. To maximize the aerodynamic efficiency at cruise, optimizations were performed using the design techniques developed internally at INTA under a research program (Programa de Termofluidodinámica). The code is a gradient-based optimizaa tion code, which uses classical finite differences approach for gradient computations. Several techniques for search direction computation are implemented for unconstrained and constrained problems. Techniques for geometry modifications are based on different approaches which include perturbation functions for the thickness and/or mean line distributions and others by Bézier curves fitting of certain degree. It is very e important to afford a real design which involves several constraints that reduce significantly the feasible design space. And the assessment of the code is needed in order to check the capabilities and the possible drawbacks. Lessons learnt will help in the development of future enhancements. In addition, the validation of the results was done using also the well-known TAU flow solver and a far-field drag method in order to determine accurately the improvement in terms of drag counts.

Aerodynamic design optimization using sensitivity analysis and computational fluid dynamics

NASA Technical Reports Server (NTRS)

Baysal, Oktay; Eleshaky, Mohamed E.

1991-01-01

A new and efficient method is presented for aerodynamic design optimization, which is based on a computational fluid dynamics (CFD)-sensitivity analysis algorithm. The method is applied to design a scramjet-afterbody configuration for an optimized axial thrust. The Euler equations are solved for the inviscid analysis of the flow, which in turn provides the objective function and the constraints. The CFD analysis is then coupled with the optimization procedure that uses a constrained minimization method. The sensitivity coefficients, i.e. gradients of the objective function and the constraints, needed for the optimization are obtained using a quasi-analytical method rather than the traditional brute force method of finite difference approximations. During the one-dimensional search of the optimization procedure, an approximate flow analysis (predicted flow) based on a first-order Taylor series expansion is used to reduce the computational cost. Finally, the sensitivity of the optimum objective function to various design parameters, which are kept constant during the optimization, is computed to predict new optimum solutions. The flow analysis of the demonstrative example are compared with the experimental data. It is shown that the method is more efficient than the traditional methods.

On the Use of CAD and Cartesian Methods for Aerodynamic Optimization

NASA Technical Reports Server (NTRS)

Nemec, M.; Aftosmis, M. J.; Pulliam, T. H.

2004-01-01

The objective for this paper is to present the development of an optimization capability for Curt3D, a Cartesian inviscid-flow analysis package. We present the construction of a new optimization framework and we focus on the following issues: 1) Component-based geometry parameterization approach using parametric-CAD models and CAPRI. A novel geometry server is introduced that addresses the issue of parallel efficiency while only sparingly consuming CAD resources; 2) The use of genetic and gradient-based algorithms for three-dimensional aerodynamic design problems. The influence of noise on the optimization methods is studied. Our goal is to create a responsive and automated framework that efficiently identifies design modifications that result in substantial performance improvements. In addition, we examine the architectural issues associated with the deployment of a CAD-based approach in a heterogeneous parallel computing environment that contains both CAD workstations and dedicated compute engines. We demonstrate the effectiveness of the framework for a design problem that features topology changes and complex geometry.

Aerodynamic Design and Numerical Analysis of Supersonic Turbine for Turbo Pump

NASA Astrophysics Data System (ADS)

Fu, Chao; Zou, Zhengping; Kong, Qingguo; Cheng, Honggui; Zhang, Weihao

2016-09-01

Supersonic turbine is widely used in the turbo pump of modern rocket. A preliminary design method for supersonic turbine has been developed considering the coupling effects of turbine and nozzle. Numerical simulation has been proceeded to validate the feasibility of the design method. As the strong shockwave reflected on the mixing plane, additional numerical simulated error would be produced by the mixing plane model in the steady CFD. So unsteady CFD is employed to investigate the aerodynamic performance of the turbine and flow field in passage. Results showed that the preliminary design method developed in this paper is suitable for designing supersonic turbine. This periodical variation of complex shockwave system influences the development of secondary flow, wake and shock-boundary layer interaction, which obviously affect the secondary loss in vane passage. The periodical variation also influences the strength of reflecting shockwave, which affects the profile loss in vane passage. Besides, high circumferential velocity at vane outlet and short blade lead to high radial pressure gradient, which makes the low kinetic energy fluid moves towards hub region and produces additional loss.

Aerothermodynamic shape optimization of hypersonic blunt bodies

NASA Astrophysics Data System (ADS)

Eyi, Sinan; Yumuşak, Mine

2015-07-01

The aim of this study is to develop a reliable and efficient design tool that can be used in hypersonic flows. The flow analysis is based on the axisymmetric Euler/Navier-Stokes and finite-rate chemical reaction equations. The equations are coupled simultaneously and solved implicitly using Newton's method. The Jacobian matrix is evaluated analytically. A gradient-based numerical optimization is used. The adjoint method is utilized for sensitivity calculations. The objective of the design is to generate a hypersonic blunt geometry that produces the minimum drag with low aerodynamic heating. Bezier curves are used for geometry parameterization. The performances of the design optimization method are demonstrated for different hypersonic flow conditions.

A Gradient-Based Multistart Algorithm for Multimodal Aerodynamic Shape Optimization Problems Based on Free-Form Deformation

NASA Astrophysics Data System (ADS)

Streuber, Gregg Mitchell

Environmental and economic factors motivate the pursuit of more fuel-efficient aircraft designs. Aerodynamic shape optimization is a powerful tool in this effort, but is hampered by the presence of multimodality in many design spaces. Gradient-based multistart optimization uses a sampling algorithm and multiple parallel optimizations to reliably apply fast gradient-based optimization to moderately multimodal problems. Ensuring that the sampled geometries remain physically realizable requires manually developing specialized linear constraints for each class of problem. Utilizing free-form deformation geometry control allows these linear constraints to be written in a geometry-independent fashion, greatly easing the process of applying the algorithm to new problems. This algorithm was used to assess the presence of multimodality when optimizing a wing in subsonic and transonic flows, under inviscid and viscous conditions, and a blended wing-body under transonic, viscous conditions. Multimodality was present in every wing case, while the blended wing-body was found to be generally unimodal.

A nonintrusive laser interferometer method for measurement of skin friction

NASA Technical Reports Server (NTRS)

Monson, D. J.

1982-01-01

A method is described for monitoring the changing thickness of a thin oil film subject to an aerodynamic shear stress using two focused laser beams. The measurement is then simply analyzed in terms of the surface skin friction of the flow. The analysis includes the effects of arbitrarily large pressure and skin friction gradients, gravity, and time varying oil temperature. It may also be applied to three dimensional flows with unknown direction. Applications are presented for a variety of flows including two dimensional flows, three dimensional swirling flows, separated flow, supersonic high Reynolds number flows, and delta wing vortical flows.

An analytical technique for approximating unsteady aerodynamics in the time domain

NASA Technical Reports Server (NTRS)

Dunn, H. J.

1980-01-01

An analytical technique is presented for approximating unsteady aerodynamic forces in the time domain. The order of elements of a matrix Pade approximation was postulated, and the resulting polynomial coefficients were determined through a combination of least squares estimates for the numerator coefficients and a constrained gradient search for the denominator coefficients which insures stable approximating functions. The number of differential equations required to represent the aerodynamic forces to a given accuracy tends to be smaller than that employed in certain existing techniques where the denominator coefficients are chosen a priori. Results are shown for an aeroelastic, cantilevered, semispan wing which indicate a good fit to the aerodynamic forces for oscillatory motion can be achieved with a matrix Pade approximation having fourth order numerator and second order denominator polynomials.

Aerodynamic Shape Optimization of Supersonic Aircraft Configurations via an Adjoint Formulation on Parallel Computers

NASA Technical Reports Server (NTRS)

Reuther, James; Alonso, Juan Jose; Rimlinger, Mark J.; Jameson, Antony

1996-01-01

This work describes the application of a control theory-based aerodynamic shape optimization method to the problem of supersonic aircraft design. The design process is greatly accelerated through the use of both control theory and a parallel implementation on distributed memory computers. Control theory is employed to derive the adjoint differential equations whose solution allows for the evaluation of design gradient information at a fraction of the computational cost required by previous design methods. The resulting problem is then implemented on parallel distributed memory architectures using a domain decomposition approach, an optimized communication schedule, and the MPI (Message Passing Interface) Standard for portability and efficiency. The final result achieves very rapid aerodynamic design based on higher order computational fluid dynamics methods (CFD). In our earlier studies, the serial implementation of this design method was shown to be effective for the optimization of airfoils, wings, wing-bodies, and complex aircraft configurations using both the potential equation and the Euler equations. In our most recent paper, the Euler method was extended to treat complete aircraft configurations via a new multiblock implementation. Furthermore, during the same conference, we also presented preliminary results demonstrating that this basic methodology could be ported to distributed memory parallel computing architectures. In this paper, our concern will be to demonstrate that the combined power of these new technologies can be used routinely in an industrial design environment by applying it to the case study of the design of typical supersonic transport configurations. A particular difficulty of this test case is posed by the propulsion/airframe integration.

Analysis of the Hessian for Aerodynamic Optimization: Inviscid Flow

NASA Technical Reports Server (NTRS)

Arian, Eyal; Ta'asan, Shlomo

1996-01-01

In this paper we analyze inviscid aerodynamic shape optimization problems governed by the full potential and the Euler equations in two and three dimensions. The analysis indicates that minimization of pressure dependent cost functions results in Hessians whose eigenvalue distributions are identical for the full potential and the Euler equations. However the optimization problems in two and three dimensions are inherently different. While the two dimensional optimization problems are well-posed the three dimensional ones are ill-posed. Oscillations in the shape up to the smallest scale allowed by the design space can develop in the direction perpendicular to the flow, implying that a regularization is required. A natural choice of such a regularization is derived. The analysis also gives an estimate of the Hessian's condition number which implies that the problems at hand are ill-conditioned. Infinite dimensional approximations for the Hessians are constructed and preconditioners for gradient based methods are derived from these approximate Hessians.

Aerodynamic Optimization of Rocket Control Surface Geometry Using Cartesian Methods and CAD Geometry

NASA Technical Reports Server (NTRS)

Nelson, Andrea; Aftosmis, Michael J.; Nemec, Marian; Pulliam, Thomas H.

2004-01-01

Aerodynamic design is an iterative process involving geometry manipulation and complex computational analysis subject to physical constraints and aerodynamic objectives. A design cycle consists of first establishing the performance of a baseline design, which is usually created with low-fidelity engineering tools, and then progressively optimizing the design to maximize its performance. Optimization techniques have evolved from relying exclusively on designer intuition and insight in traditional trial and error methods, to sophisticated local and global search methods. Recent attempts at automating the search through a large design space with formal optimization methods include both database driven and direct evaluation schemes. Databases are being used in conjunction with surrogate and neural network models as a basis on which to run optimization algorithms. Optimization algorithms are also being driven by the direct evaluation of objectives and constraints using high-fidelity simulations. Surrogate methods use data points obtained from simulations, and possibly gradients evaluated at the data points, to create mathematical approximations of a database. Neural network models work in a similar fashion, using a number of high-fidelity database calculations as training iterations to create a database model. Optimal designs are obtained by coupling an optimization algorithm to the database model. Evaluation of the current best design then gives either a new local optima and/or increases the fidelity of the approximation model for the next iteration. Surrogate methods have also been developed that iterate on the selection of data points to decrease the uncertainty of the approximation model prior to searching for an optimal design. The database approximation models for each of these cases, however, become computationally expensive with increase in dimensionality. Thus the method of using optimization algorithms to search a database model becomes problematic as the number of design variables is increased.

Drag De-Orbit Device (D3): A Retractable Device for CubeSat Attitude and Orbit Control using Aerodynamic Forces

NASA Technical Reports Server (NTRS)

Guglielmo, David; Omar, Sanny R.; Bevilacqua, Riccardo

2017-01-01

The increasing number of CubeSats being launched has raised concerns about orbital debris since most of these satellites have no means of active orbit control. Some technologies exist to increase the surface area of a CubeSat and expedite de-orbit due to aerodynamic drag in low Earth orbit, but most of these devices cannot be retracted and hence cannot be used for orbital maneuvering. This paper discusses the De-Orbit Drag Device (D3) module that is capable of de-orbiting a 12U, 15kg CubeSat from a 700 km circular orbit in under 25 years and can be deployed and retracted to modulate the aerodynamic drag force experienced by the satellite. This facilitates orbital maneuvering using aerodynamic drag and the active targeting of a de-orbit location. In addition, the geometry of this drag device provides 3-axis attitude stabilization of the host CubeSat using aerodynamic and gravity gradient torques which is useful for many missions and provides a predictable aerodynamic profile for use in orbital maneuvering algorithms.

A numerical and theoretical study on the aerodynamics of a rhinoceros beetle (Trypoxlyus dichotomus) and optimization of its wing kinematics in hover

NASA Astrophysics Data System (ADS)

Oh, Sehyeong; Lee, Boogeon; Park, Hyungmin; Choi, Haecheon

2017-11-01

We investigate a hovering rhinoceros beetle using numerical simulation and blade element theory. Numerical simulations are performed using an immersed boundary method. In the simulation, the hindwings are modeled as a rigid flat plate, and three-dimensionally scanned elytra and body are used. The results of simulation indicate that the lift force generated by the hindwings alone is sufficient to support the weight, and the elytra generate negligible lift force. Considering the hindwings only, we present a blade element model based on quasi-steady assumptions to identify the mechanisms of aerodynamic force generation and power expenditure in the hovering flight of a rhinoceros beetle. We show that the results from the present blade element model are in excellent agreement with numerical ones. Based on the current blade element model, we find the optimal wing kinematics minimizing the aerodynamic power requirement using a hybrid optimization algorithm combining a clustering genetic algorithm with a gradient-based optimizer. We show that the optimal wing kinematics reduce the aerodynamic power consumption, generating enough lift force to support the weight. This research was supported by a Grant to Bio-Mimetic Robot Research Center Funded by Defense Acquisition Program Administration, and by Agency for Defense Development (UD130070ID) and NRF-2016R1E1A1A02921549 of the MSIP of Korea.

Constrained Multipoint Aerodynamic Shape Optimization Using an Adjoint Formulation and Parallel Computers

NASA Technical Reports Server (NTRS)

Reuther, James; Jameson, Antony; Alonso, Juan Jose; Rimlinger, Mark J.; Saunders, David

1997-01-01

An aerodynamic shape optimization method that treats the design of complex aircraft configurations subject to high fidelity computational fluid dynamics (CFD), geometric constraints and multiple design points is described. The design process will be greatly accelerated through the use of both control theory and distributed memory computer architectures. Control theory is employed to derive the adjoint differential equations whose solution allows for the evaluation of design gradient information at a fraction of the computational cost required by previous design methods. The resulting problem is implemented on parallel distributed memory architectures using a domain decomposition approach, an optimized communication schedule, and the MPI (Message Passing Interface) standard for portability and efficiency. The final result achieves very rapid aerodynamic design based on a higher order CFD method. In order to facilitate the integration of these high fidelity CFD approaches into future multi-disciplinary optimization (NW) applications, new methods must be developed which are capable of simultaneously addressing complex geometries, multiple objective functions, and geometric design constraints. In our earlier studies, we coupled the adjoint based design formulations with unconstrained optimization algorithms and showed that the approach was effective for the aerodynamic design of airfoils, wings, wing-bodies, and complex aircraft configurations. In many of the results presented in these earlier works, geometric constraints were satisfied either by a projection into feasible space or by posing the design space parameterization such that it automatically satisfied constraints. Furthermore, with the exception of reference 9 where the second author initially explored the use of multipoint design in conjunction with adjoint formulations, our earlier works have focused on single point design efforts. Here we demonstrate that the same methodology may be extended to treat complete configuration designs subject to multiple design points and geometric constraints. Examples are presented for both transonic and supersonic configurations ranging from wing alone designs to complex configuration designs involving wing, fuselage, nacelles and pylons.

Investigation of rotor blade tip-vortex aerodynamics

NASA Technical Reports Server (NTRS)

Lewellen, W. S.

1971-01-01

Several aspects of the aerodynamics of rotor blade tip vortices are examined. Two particular categories are dealt with; (1) dynamic loads on a blade passing close to or intersecting a trailing vortex, and (2) the response of the trailing vortex core to changes in the flow. Results for both categories are in reasonable agreement with existing data, although lower pressure gradients were obtained than anticipated for category one. A correlation between trailing edge sweep angle at the tip and vortex core size was noted for category two.

A nonintrusive laser interferometer method for measurement of skin friction

NASA Technical Reports Server (NTRS)

Monson, D. J.

1983-01-01

A method is described for monitoring the changing thickness of a thin oil film subject to an aerodynamic shear stress using two focused laser beams. The measurement is then simply analyzed in terms of the surface skin friction of the flow. The analysis includes the effects of arbitrarily large pressure and skin friction gradients, gravity, and time varying oil temperature. It may also be applied to three dimensional flows with unknown direction. Applications are presented for a variety of flows, including two dimensional flows, three dimensional swirling flows, separated flow, supersonic high Reynolds number flows, and delta wing vortical flows. Previously announced in STAR as N83-12393

Low Complexity Models to improve Incomplete Sensitivities for Shape Optimization

NASA Astrophysics Data System (ADS)

Stanciu, Mugurel; Mohammadi, Bijan; Moreau, Stéphane

2003-01-01

The present global platform for simulation and design of multi-model configurations treat shape optimization problems in aerodynamics. Flow solvers are coupled with optimization algorithms based on CAD-free and CAD-connected frameworks. Newton methods together with incomplete expressions of gradients are used. Such incomplete sensitivities are improved using reduced models based on physical assumptions. The validity and the application of this approach in real-life problems are presented. The numerical examples concern shape optimization for an airfoil, a business jet and a car engine cooling axial fan.

Sensitivity analysis, approximate analysis, and design optimization for internal and external viscous flows

NASA Technical Reports Server (NTRS)

Taylor, Arthur C., III; Hou, Gene W.; Korivi, Vamshi M.

1991-01-01

A gradient-based design optimization strategy for practical aerodynamic design applications is presented, which uses the 2D thin-layer Navier-Stokes equations. The strategy is based on the classic idea of constructing different modules for performing the major tasks such as function evaluation, function approximation and sensitivity analysis, mesh regeneration, and grid sensitivity analysis, all driven and controlled by a general-purpose design optimization program. The accuracy of aerodynamic shape sensitivity derivatives is validated on two viscous test problems: internal flow through a double-throat nozzle and external flow over a NACA 4-digit airfoil. A significant improvement in aerodynamic performance has been achieved in both cases. Particular attention is given to a consistent treatment of the boundary conditions in the calculation of the aerodynamic sensitivity derivatives for the classic problems of external flow over an isolated lifting airfoil on 'C' or 'O' meshes.

Aerodynamic Shape Optimization of Supersonic Aircraft Configurations via an Adjoint Formulation on Parallel Computers

NASA Technical Reports Server (NTRS)

Reuther, James; Alonso, Juan Jose; Rimlinger, Mark J.; Jameson, Antony

1996-01-01

This work describes the application of a control theory-based aerodynamic shape optimization method to the problem of supersonic aircraft design. The design process is greatly accelerated through the use of both control theory and a parallel implementation on distributed memory computers. Control theory is employed to derive the adjoint differential equations whose solution allows for the evaluation of design gradient information at a fraction of the computational cost required by previous design methods (13, 12, 44, 38). The resulting problem is then implemented on parallel distributed memory architectures using a domain decomposition approach, an optimized communication schedule, and the MPI (Message Passing Interface) Standard for portability and efficiency. The final result achieves very rapid aerodynamic design based on higher order computational fluid dynamics methods (CFD). In our earlier studies, the serial implementation of this design method (19, 20, 21, 23, 39, 25, 40, 41, 42, 43, 9) was shown to be effective for the optimization of airfoils, wings, wing-bodies, and complex aircraft configurations using both the potential equation and the Euler equations (39, 25). In our most recent paper, the Euler method was extended to treat complete aircraft configurations via a new multiblock implementation. Furthermore, during the same conference, we also presented preliminary results demonstrating that the basic methodology could be ported to distributed memory parallel computing architectures [241. In this paper, our concem will be to demonstrate that the combined power of these new technologies can be used routinely in an industrial design environment by applying it to the case study of the design of typical supersonic transport configurations. A particular difficulty of this test case is posed by the propulsion/airframe integration.

Development of an Innovative Algorithm for Aerodynamics-Structure Interaction Using Lattice Boltzmann Method

NASA Technical Reports Server (NTRS)

Mei, Ren-Wei; Shyy, Wei; Yu, Da-Zhi; Luo, Li-Shi; Rudy, David (Technical Monitor)

2001-01-01

The lattice Boltzmann equation (LBE) is a kinetic formulation which offers an alternative computational method capable of solving fluid dynamics for various systems. Major advantages of the method are owing to the fact that the solution for the particle distribution functions is explicit, easy to implement, and the algorithm is natural to parallelize. In this final report, we summarize the works accomplished in the past three years. Since most works have been published, the technical details can be found in the literature. Brief summary will be provided in this report. In this project, a second-order accurate treatment of boundary condition in the LBE method is developed for a curved boundary and tested successfully in various 2-D and 3-D configurations. To evaluate the aerodynamic force on a body in the context of LBE method, several force evaluation schemes have been investigated. A simple momentum exchange method is shown to give reliable and accurate values for the force on a body in both 2-D and 3-D cases. Various 3-D LBE models have been assessed in terms of efficiency, accuracy, and robustness. In general, accurate 3-D results can be obtained using LBE methods. The 3-D 19-bit model is found to be the best one among the 15-bit, 19-bit, and 27-bit LBE models. To achieve desired grid resolution and to accommodate the far field boundary conditions in aerodynamics computations, a multi-block LBE method is developed by dividing the flow field into various blocks each having constant lattice spacing. Substantial contribution to the LBE method is also made through the development of a new, generalized lattice Boltzmann equation constructed in the moment space in order to improve the computational stability, detailed theoretical analysis on the stability, dispersion, and dissipation characteristics of the LBE method, and computational studies of high Reynolds number flows with singular gradients. Finally, a finite difference-based lattice Boltzmann method is developed for inviscid compressible flows.

Electric-car simulation

NASA Technical Reports Server (NTRS)

Chapman, C. P.; Slusser, R. A.

1980-01-01

PARAMET, interactive simulation program for parametric studies of electric vehicles, guides user through simulation by menu and series of prompts for input parameters. Program considers aerodynamic drag, rolling resistance, linear and rotational acceleration, and road gradient as forces acting on vehicle.

Numerical and Experimental Validation of the Optimization Methodologies for a Wing-Tip Structure Equipped with Conventional and Morphing Ailerons =

NASA Astrophysics Data System (ADS)

Koreanschi, Andreea

In order to answer the problem of 'how to reduce the aerospace industry's environment footprint?' new morphing technologies were developed. These technologies were aimed at reducing the aircraft's fuel consumption through reduction of the wing drag. The morphing concept used in the present research consists of replacing the conventional aluminium upper surface of the wing with a flexible composite skin for morphing abilities. For the ATR-42 'Morphing wing' project, the wing models were manufactured entirely from composite materials and the morphing region was optimized for flexibility. In this project two rigid wing models and an active morphing wing model were designed, manufactured and wind tunnel tested. For the CRIAQ MDO 505 project, a full scale wing-tip equipped with two types of ailerons, conventional and morphing, was designed, optimized, manufactured, bench and wind tunnel tested. The morphing concept was applied on a real wing internal structure and incorporated aerodynamic, structural and control constraints specific to a multidisciplinary approach. Numerical optimization, aerodynamic analysis and experimental validation were performed for both the CRIAQ MDO 505 full scale wing-tip demonstrator and the ATR-42 reduced scale wing models. In order to improve the aerodynamic performances of the ATR-42 and CRIAQ MDO 505 wing airfoils, three global optimization algorithms were developed, tested and compared. The three algorithms were: the genetic algorithm, the artificial bee colony and the gradient descent. The algorithms were coupled with the two-dimensional aerodynamic solver XFoil. XFoil is known for its rapid convergence, robustness and use of the semi-empirical e n method for determining the position of the flow transition from laminar to turbulent. Based on the performance comparison between the algorithms, the genetic algorithm was chosen for the optimization of the ATR-42 and CRIAQ MDO 505 wing airfoils. The optimization algorithm was improved during the CRIAQ MDO 505 project for convergence speed by introducing a two-step cross-over function. Structural constraints were introduced in the algorithm at each aero-structural optimization interaction, allowing a better manipulation of the algorithm and giving it more capabilities of morphing combinations. The CRIAQ MDO 505 project envisioned a morphing aileron concept for the morphing upper surface wing. For this morphing aileron concept, two optimization methods were developed. The methods used the already developed genetic algorithm and each method had a different design concept. The first method was based on the morphing upper surface concept, using actuation points to achieve the desired shape. The second method was based on the hinge rotation concept of the conventional aileron but applied at multiple nodes along the aileron camber to achieve the desired shape. Both methods were constrained by manufacturing and aerodynamic requirements. The purpose of the morphing aileron methods was to obtain an aileron shape with a smoother pressure distribution gradient during deflection than the conventional aileron. The aerodynamic optimization results were used for the structural optimization and design of the wing, particularly the flexible composite skin. Due to the structural changes performed on the initial wing-tip structure, an aeroelastic behaviour analysis, more specific on flutter phenomenon, was performed. The analyses were done to ensure the structural integrity of the wing-tip demonstrator during wind tunnel tests. Three wind tunnel tests were performed for the CRIAQ MDO 505 wing-tip demonstrator at the IAR-NRC subsonic wind tunnel facility in Ottawa. The first two tests were performed for the wing-tip equipped with conventional aileron. The purpose of these tests was to validate the control system designed for the morphing upper surface, the numerical optimization and aerodynamic analysis and to evaluate the optimization efficiency on the boundary layer behaviour and the wing drag. The third set of wind tunnel tests was performed on the wing-tip equipped with a morphing aileron. The purpose of this test was to evaluate the performances of the morphing aileron, in conjunction with the active morphing upper surface, and their effect on the lift, drag and boundary layer behaviour. Transition data, obtained from Infrared Thermography, and pressure data, extracted from Kulite and pressure taps recordings, were used to validate the numerical optimization and aerodynamic performances of the wing-tip demonstrator. A set of wind tunnel tests was performed on the ATR-42 rigid wing models at the Price-Paidoussis subsonic wind tunnel at Ecole de technologie Superieure. The results from the pressure taps recordings were used to validate the numerical optimization. A second derivative of the pressure distribution method was applied to evaluate the transition region on the upper surface of the wing models for comparison with the numerical transition values. (Abstract shortened by ProQuest.).

Building complex simulations rapidly using MATRIX(x): The Space Station redesign

NASA Technical Reports Server (NTRS)

Carrington, C. K.

1994-01-01

MSFC's quick response to the Space Station redesign effort last year required the development of a computer simulation to model the attitude and station-keeping dynamics of a complex body with rotating solar arrays in orbit around the Earth. The simulation was written using a rapid-prototyping graphical simulation and design tool called MATRIX(x) and provided the capability to quickly remodel complex configuration changes by icon manipulation using a mouse. The simulation determines time-dependent inertia properties, and models forces and torques from gravity-gradient, solar radiation, and aerodynamic disturbances. Surface models are easily built from a selection of beams, plates, tetrahedrons, and cylinders. An optimization scheme was written to determine the torque equilibrium attitudes that balance gravity-gradient and aerodynamic torques over an orbit, and propellant-usage estimates were determined. The simulation has been adapted to model the attitude dynamics for small spacecraft.

Comparisons of several aerodynamic methods for application to dynamic loads analyses

NASA Technical Reports Server (NTRS)

Kroll, R. I.; Miller, R. D.

1976-01-01

The results of a study are presented in which the applicability at subsonic speeds of several aerodynamic methods for predicting dynamic gust loads on aircraft, including active control systems, was examined and compared. These aerodynamic methods varied from steady state to an advanced unsteady aerodynamic formulation. Brief descriptions of the structural and aerodynamic representations and of the motion and load equations are presented. Comparisons of numerical results achieved using the various aerodynamic methods are shown in detail. From these results, aerodynamic representations for dynamic gust analyses are identified. It was concluded that several aerodynamic methods are satisfactory for dynamic gust analyses of configurations having either controls fixed or active control systems that primarily affect the low frequency rigid body aircraft response.

Global Design Optimization for Aerodynamics and Rocket Propulsion Components

NASA Technical Reports Server (NTRS)

Shyy, Wei; Papila, Nilay; Vaidyanathan, Rajkumar; Tucker, Kevin; Turner, James E. (Technical Monitor)

2000-01-01

Modern computational and experimental tools for aerodynamics and propulsion applications have matured to a stage where they can provide substantial insight into engineering processes involving fluid flows, and can be fruitfully utilized to help improve the design of practical devices. In particular, rapid and continuous development in aerospace engineering demands that new design concepts be regularly proposed to meet goals for increased performance, robustness and safety while concurrently decreasing cost. To date, the majority of the effort in design optimization of fluid dynamics has relied on gradient-based search algorithms. Global optimization methods can utilize the information collected from various sources and by different tools. These methods offer multi-criterion optimization, handle the existence of multiple design points and trade-offs via insight into the entire design space, can easily perform tasks in parallel, and are often effective in filtering the noise intrinsic to numerical and experimental data. However, a successful application of the global optimization method needs to address issues related to data requirements with an increase in the number of design variables, and methods for predicting the model performance. In this article, we review recent progress made in establishing suitable global optimization techniques employing neural network and polynomial-based response surface methodologies. Issues addressed include techniques for construction of the response surface, design of experiment techniques for supplying information in an economical manner, optimization procedures and multi-level techniques, and assessment of relative performance between polynomials and neural networks. Examples drawn from wing aerodynamics, turbulent diffuser flows, gas-gas injectors, and supersonic turbines are employed to help demonstrate the issues involved in an engineering design context. Both the usefulness of the existing knowledge to aid current design practices and the need for future research are identified.

Multi-disciplinary optimization of aeroservoelastic systems

NASA Technical Reports Server (NTRS)

Karpel, Mordechay

1990-01-01

Efficient analytical and computational tools for simultaneous optimal design of the structural and control components of aeroservoelastic systems are presented. The optimization objective is to achieve aircraft performance requirements and sufficient flutter and control stability margins with a minimal weight penalty and without violating the design constraints. Analytical sensitivity derivatives facilitate an efficient optimization process which allows a relatively large number of design variables. Standard finite element and unsteady aerodynamic routines are used to construct a modal data base. Minimum State aerodynamic approximations and dynamic residualization methods are used to construct a high accuracy, low order aeroservoelastic model. Sensitivity derivatives of flutter dynamic pressure, control stability margins and control effectiveness with respect to structural and control design variables are presented. The performance requirements are utilized by equality constraints which affect the sensitivity derivatives. A gradient-based optimization algorithm is used to minimize an overall cost function. A realistic numerical example of a composite wing with four controls is used to demonstrate the modeling technique, the optimization process, and their accuracy and efficiency.

Multidisciplinary optimization of aeroservoelastic systems using reduced-size models

NASA Technical Reports Server (NTRS)

Karpel, Mordechay

1992-01-01

Efficient analytical and computational tools for simultaneous optimal design of the structural and control components of aeroservoelastic systems are presented. The optimization objective is to achieve aircraft performance requirements and sufficient flutter and control stability margins with a minimal weight penalty and without violating the design constraints. Analytical sensitivity derivatives facilitate an efficient optimization process which allows a relatively large number of design variables. Standard finite element and unsteady aerodynamic routines are used to construct a modal data base. Minimum State aerodynamic approximations and dynamic residualization methods are used to construct a high accuracy, low order aeroservoelastic model. Sensitivity derivatives of flutter dynamic pressure, control stability margins and control effectiveness with respect to structural and control design variables are presented. The performance requirements are utilized by equality constraints which affect the sensitivity derivatives. A gradient-based optimization algorithm is used to minimize an overall cost function. A realistic numerical example of a composite wing with four controls is used to demonstrate the modeling technique, the optimization process, and their accuracy and efficiency.

A method of self-pursued boundary value on a body and the Magnus effect calculated with this method

NASA Astrophysics Data System (ADS)

Yoshino, Fumio; Hayashi, Tatsuo; Waka, Ryoji

1991-03-01

A computational method, designated 'SPB', is proposed for the automatic determination of the stream function Phi on an arbitrarily profiled body without recourse to empirical factors. The method is applied to the case of a rotating, circular cross-section cylinder in a uniform shear flow, and the results obtained are compared with those of both the method in which the value of Phi is fixed on a body and the conventional empirical method; it is in view of this established that the SPB method is very efficient and applicable to both steady and unsteady flows. The SPB method, in addition to yielding the aerodynamic forces acting on a cylinder, shows that the Magnus effect lift force decreases as the velocity gradient of the shear flow increases while the cylinder's rotational speed is kept constant.

Multi-fidelity and multi-disciplinary design optimization of supersonic business jets

NASA Astrophysics Data System (ADS)

Choi, Seongim

Supersonic jets have been drawing great attention after the end of service for the Concorde was announced on April of 2003. It is believed, however, that civilian supersonic aircraft may make a viable return in the business jet market. This thesis focuses on the design optimization of feasible supersonic business jet configurations. Preliminary design techniques for mitigation of ground sonic boom are investigated while ensuring that all relevant disciplinary constraints are satisfied (including aerodynamic performance, propulsion, stability & control and structures.) In order to achieve reasonable confidence in the resulting designs, high-fidelity simulations are required, making the entire design process both expensive and complex. In order to minimize the computational cost, surrogate/approximate models are constructed using a hierarchy of different fidelity analysis tools including PASS, A502/Panair and Euler/NS codes. Direct search methods such as Genetic Algorithms (GAs) and a nonlinear SIMPLEX are employed to designs in searches of large and noisy design spaces. A local gradient-based search method can be combined with these global search methods for small modifications of candidate optimum designs. The Mesh Adaptive Direct Search (MADS) method can also be used to explore the design space using a solution-adaptive grid refinement approach. These hybrid approaches, both in search methodology and surrogate model construction, are shown to result in designs with reductions in sonic boom and improved aerodynamic performance.

Atmospheric stability and diurnal patterns of aeolian saltation on the Llano Estacado

USDA-ARS?s Scientific Manuscript database

Aeolian transport is driven by aerodynamic surface stress imposed by turbulent winds in the Earth’s atmospheric boundary layer (ABL). ABL regime is influenced by stratification, which can either enhance or suppress production of turbulence by shear associated with the vertical gradient of streamwise...

Tether fundamentals

NASA Technical Reports Server (NTRS)

Carroll, J. A.

1986-01-01

Some fundamental aspects of tethers are presented and briefly discussed. The effects of gravity gradients, dumbbell libration in circular orbits, tether control strategies and impact hazards for tethers are among those fundamentals. Also considered are aerodynamic drag, constraints in momentum transfer applications and constraints with permanently deployed tethers. The theoretical feasibility of these concepts are reviewed.

Energy Balance Bowen Ratio (EBBR) Handbook

DOE Office of Scientific and Technical Information (OSTI.GOV)

Cook, D. R.

2016-01-01

The Energy Balance Bowen Ratio (EBBR) system produces 30-minute estimates of the vertical fluxes of sensible and latent heat at the local surface. Flux estimates are calculated from observations of net radiation, soil surface heat flux, and the vertical gradients of temperature and relative humidity (RH). Meteorological data collected by the EBBR are used to calculate bulk aerodynamic fluxes, which are used in the Bulk Aerodynamic Technique (BA) EBBR value-added product (VAP) to replace sunrise and sunset spikes in the flux data. A unique aspect of the system is the automatic exchange mechanism (AEM), which helps to reduce errors frommore » instrument offset drift.« less

Energy Balance Bowen Ratio Station (EBBR) Handbook

DOE Office of Scientific and Technical Information (OSTI.GOV)

Cook, DR

2011-02-23

The energy balance Bowen ratio (EBBR) system produces 30-minute estimates of the vertical fluxes of sensible and latent heat at the local surface. Flux estimates are calculated from observations of net radiation, soil surface heat flux, and the vertical gradients of temperature and relative humidity (RH). Meteorological data collected by the EBBR are used to calculate bulk aerodynamic fluxes, which are used in the Bulk Aerodynamic Technique (BA) EBBR value-added product (VAP) to replace sunrise and sunset spikes in the flux data. A unique aspect of the system is the automatic exchange mechanism (AEM), which helps to reduce errors frommore » instrument offset drift.« less

Aerodynamic Design on Unstructured Grids for Turbulent Flows

NASA Technical Reports Server (NTRS)

Anderson, W. Kyle; Bonhaus, Daryl L.

1997-01-01

An aerodynamic design algorithm for turbulent flows using unstructured grids is described. The current approach uses adjoint (costate) variables for obtaining derivatives of the cost function. The solution of the adjoint equations is obtained using an implicit formulation in which the turbulence model is fully coupled with the flow equations when solving for the costate variables. The accuracy of the derivatives is demonstrated by comparison with finite-difference gradients and a few example computations are shown. In addition, a user interface is described which significantly reduces the time required for setting up the design problems. Recommendations on directions of further research into the Navier Stokes design process are made.

Economical Unsteady High-Fidelity Aerodynamics for Structural Optimization with a Flutter Constraint

NASA Technical Reports Server (NTRS)

Bartels, Robert E.; Stanford, Bret K.

2017-01-01

Structural optimization with a flutter constraint for a vehicle designed to fly in the transonic regime is a particularly difficult task. In this speed range, the flutter boundary is very sensitive to aerodynamic nonlinearities, typically requiring high-fidelity Navier-Stokes simulations. However, the repeated application of unsteady computational fluid dynamics to guide an aeroelastic optimization process is very computationally expensive. This expense has motivated the development of methods that incorporate aspects of the aerodynamic nonlinearity, classical tools of flutter analysis, and more recent methods of optimization. While it is possible to use doublet lattice method aerodynamics, this paper focuses on the use of an unsteady high-fidelity aerodynamic reduced order model combined with successive transformations that allows for an economical way of utilizing high-fidelity aerodynamics in the optimization process. This approach is applied to the common research model wing structural design. As might be expected, the high-fidelity aerodynamics produces a heavier wing than that optimized with doublet lattice aerodynamics. It is found that the optimized lower skin of the wing using high-fidelity aerodynamics differs significantly from that using doublet lattice aerodynamics.

Aerodynamics of a linear oscillating cascade

NASA Technical Reports Server (NTRS)

Buffum, Daniel H.; Fleeter, Sanford

1990-01-01

The steady and unsteady aerodynamics of a linear oscillating cascade are investigated using experimental and computational methods. Experiments are performed to quantify the torsion mode oscillating cascade aerodynamics of the NASA Lewis Transonic Oscillating Cascade for subsonic inlet flowfields using two methods: simultaneous oscillation of all the cascaded airfoils at various values of interblade phase angle, and the unsteady aerodynamic influence coefficient technique. Analysis of these data and correlation with classical linearized unsteady aerodynamic analysis predictions indicate that the wind tunnel walls enclosing the cascade have, in some cases, a detrimental effect on the cascade unsteady aerodynamics. An Euler code for oscillating cascade aerodynamics is modified to incorporate improved upstream and downstream boundary conditions and also the unsteady aerodynamic influence coefficient technique. The new boundary conditions are shown to improve the unsteady aerodynamic influence coefficient technique. The new boundary conditions are shown to improve the unsteady aerodynamic predictions of the code, and the computational unsteady aerodynamic influence coefficient technique is shown to be a viable alternative for calculation of oscillating cascade aerodynamics.

Airfoil optimization for unsteady flows with application to high-lift noise reduction

NASA Astrophysics Data System (ADS)

Rumpfkeil, Markus Peer

The use of steady-state aerodynamic optimization methods in the computational fluid dynamic (CFD) community is fairly well established. In particular, the use of adjoint methods has proven to be very beneficial because their cost is independent of the number of design variables. The application of numerical optimization to airframe-generated noise, however, has not received as much attention, but with the significant quieting of modern engines, airframe noise now competes with engine noise. Optimal control techniques for unsteady flows are needed in order to be able to reduce airframe-generated noise. In this thesis, a general framework is formulated to calculate the gradient of a cost function in a nonlinear unsteady flow environment via the discrete adjoint method. The unsteady optimization algorithm developed in this work utilizes a Newton-Krylov approach since the gradient-based optimizer uses the quasi-Newton method BFGS, Newton's method is applied to the nonlinear flow problem, GMRES is used to solve the resulting linear problem inexactly, and last but not least the linear adjoint problem is solved using Bi-CGSTAB. The flow is governed by the unsteady two-dimensional compressible Navier-Stokes equations in conjunction with a one-equation turbulence model, which are discretized using structured grids and a finite difference approach. The effectiveness of the unsteady optimization algorithm is demonstrated by applying it to several problems of interest including shocktubes, pulses in converging-diverging nozzles, rotating cylinders, transonic buffeting, and an unsteady trailing-edge flow. In order to address radiated far-field noise, an acoustic wave propagation program based on the Ffowcs Williams and Hawkings (FW-H) formulation is implemented and validated. The general framework is then used to derive the adjoint equations for a novel hybrid URANS/FW-H optimization algorithm in order to be able to optimize the shape of airfoils based on their calculated far-field pressure fluctuations. Validation and application results for this novel hybrid URANS/FW-H optimization algorithm show that it is possible to optimize the shape of an airfoil in an unsteady flow environment to minimize its radiated far-field noise while maintaining good aerodynamic performance.

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 calculated the generalized aerodynamic forces, eigenvalues, and response amplitudes. The eigenvalues determine the flutter frequency and damping. As a test case, the flutter of a helical fan was calculated with LINFLUX and compared with calculations from TURBO-AE, a nonlinear time domain code, and from ASTROP2, a code based on linear unsteady aerodynamics.

Tactical missile aerodynamics

NASA Technical Reports Server (NTRS)

Hemsch, Michael J. (Editor); Nielsen, Jack N. (Editor)

1986-01-01

The present conference on tactical missile aerodynamics discusses autopilot-related aerodynamic design considerations, flow visualization methods' role in the study of high angle-of-attack aerodynamics, low aspect ratio wing behavior at high angle-of-attack, supersonic airbreathing propulsion system inlet design, missile bodies with noncircular cross section and bank-to-turn maneuvering capabilities, 'waverider' supersonic cruise missile concepts and design methods, asymmetric vortex sheding phenomena from bodies-of-revolution, and swept shock wave/boundary layer interaction phenomena. Also discussed are the assessment of aerodynamic drag in tactical missiles, the analysis of supersonic missile aerodynamic heating, the 'equivalent angle-of-attack' concept for engineering analysis, the vortex cloud model for body vortex shedding and tracking, paneling methods with vorticity effects and corrections for nonlinear compressibility, the application of supersonic full potential method to missile bodies, Euler space marching methods for missiles, three-dimensional missile boundary layers, and an analysis of exhaust plumes and their interaction with missile airframes.

An Influence Function Method for Predicting Store Aerodynamic Characteristics during Weapon Separation,

DTIC Science & Technology

1981-05-14

8217 AO-Ail 777 GRUMMAN AEROSPACE CORP BETHPAGE NY F/G 20/4 AN INFLUENCE FUNCTION METHOD FOR PREDICTING STORE AERODYNAMIC C--ETCCU) MAY 8 1 R MEYER, A...CENKO, S YARDS UNCLASSIFIED N ’.**~~N**n I EHEEKI j~j .25 Q~4 111110 111_L 5. AN INFLUENCE FUNCTION METHOD FOR PREDICTING STORE AERODYNAMIC...extended to their logical conclusion one is led quite naturally to consideration of an " Influence Function Method" for I predicting store aerodynamic

Assessment of fluctuating pressure gradient using acceleration spectra in near wall flows

NASA Astrophysics Data System (ADS)

Cadel, Daniel; Lowe, K. Todd

2015-11-01

Separation of contributions to the fluctuating acceleration from pressure gradient fluctuations and viscous shear fluctuations in the frequency domain is examined in a turbulent boundary layer. Past work leveraging turbulent accelerations for pressure gradient measurements has neglected the viscous shear term from the momentum equation--an invalid assumption in the case of near wall flows. The present study seeks to account for the influence of the viscous shear term and spectrally reject its contribution, which is thought to be concentrated at higher frequencies. Spectra of velocity and acceleration fluctuations in a flat plate, zero pressure gradient turbulent boundary layer at a momentum thickness Reynolds number of 7500 are measured using a spatially resolving three-component laser Doppler velocimeter. This canonical case data is applied for validation of the spectral approach for future application in more complex aerodynamic flows.

Practical Applications of a Building Method to Construct Aerodynamic Database of Guided Missile Using Wind Tunnel Test Data

NASA Astrophysics Data System (ADS)

Kim, Duk-hyun; Lee, Hyoung-Jin

2018-04-01

A study of efficient aerodynamic database modeling method was conducted. A creation of database using periodicity and symmetry characteristic of missile aerodynamic coefficient was investigated to minimize the number of wind tunnel test cases. In addition, studies of how to generate the aerodynamic database when the periodicity changes due to installation of protuberance and how to conduct a zero calibration were carried out. Depending on missile configurations, the required number of test cases changes and there exist tests that can be omitted. A database of aerodynamic on deflection angle of control surface can be constituted using phase shift. A validity of modeling method was demonstrated by confirming that the result which the aerodynamic coefficient calculated by using the modeling method was in agreement with wind tunnel test results.

Evolutionary Optimization of Centrifugal Nozzles for Organic Vapours

NASA Astrophysics Data System (ADS)

Persico, Giacomo

2017-03-01

This paper discusses the shape-optimization of non-conventional centrifugal turbine nozzles for Organic Rankine Cycle applications. The optimal aerodynamic design is supported by the use of a non-intrusive, gradient-free technique specifically developed for shape optimization of turbomachinery profiles. The method is constructed as a combination of a geometrical parametrization technique based on B-Splines, a high-fidelity and experimentally validated Computational Fluid Dynamic solver, and a surrogate-based evolutionary algorithm. The non-ideal gas behaviour featuring the flow of organic fluids in the cascades of interest is introduced via a look-up-table approach, which is rigorously applied throughout the whole optimization process. Two transonic centrifugal nozzles are considered, featuring very different loading and radial extension. The use of a systematic and automatic design method to such a non-conventional configuration highlights the character of centrifugal cascades; the blades require a specific and non-trivial definition of the shape, especially in the rear part, to avoid the onset of shock waves. It is shown that the optimization acts in similar way for the two cascades, identifying an optimal curvature of the blade that both provides a relevant increase of cascade performance and a reduction of downstream gradients.

Thermal and ghost reflection modeling for a 180-deg. field-of-view long-wave infrared lens

NASA Astrophysics Data System (ADS)

Shi, Weimin; Couture, Michael E.

2001-03-01

Optics 1, Inc. has successfully designed and developed a 180 degree(s) field of view long wave infrared lens for USAF/AFRL under SBIR phase I and II funded projects in support of the multi-national Programmable Integrated Ordinance Suite (PIOS) program. In this paper, a procedure is presented on how to evaluate image degradation caused by asymmetric aerodynamic dome heating. In addition, a thermal gradient model is proposed to evaluate degradation caused by axial temperature gradient throughout the entire PIOS lens. Finally, a ghost reflection analysis is demonstrated with non-sequential model.

Guidebook for analysis of tether applications

NASA Technical Reports Server (NTRS)

Carroll, J. A.

1985-01-01

This guidebook is intended as a tool to facilitate initial analyses of proposed tether applications in space. Topics disscussed include: orbit and orbit transfer equations; orbital perturbations; aerodynamic drag; thermal balance; micrometeoroids; gravity gradient effects; tether control strategies; momentum transfer; orbit transfer by tethered release/rendezvous; impact hazards for tethers; electrodynamic tether principles; and electrodynamic libration control issues.

Magsat attitude dynamics and control: Some observations and explanations

NASA Technical Reports Server (NTRS)

Stengle, T. H.

1980-01-01

Before its reentry 7 months after launch, Magsat transmitted an abundance of valuable data for mapping the Earth's magnetic field. As an added benefit, a wealth of attitude data for study by spacecraft dynamicists was also collected. Because of its unique configuration, Magsat presented new control problems. With its aerodynamic trim boom, attitude control was given an added dimension. Minimization of attitude drift, which could be mapped in relative detail, became the goal. Momentum control, which was accomplished by pitching the spacecraft in order to balance aerodynamic and gravity gradient torques, was seldom difficult to achieve. Several interesting phenomena observed as part of this activity included occasional momentum wheel instability and a rough correlation between solar flux and the pitch angle required to maintain acceptable momentum. An overview is presented of the attitude behavior of Magsat and some of the control problems encountered. Plausible explanations for some of this behavior are offered. Some of the control philosophy used during the mission is examined and aerodynamic trimming operations are summarized.

Acoustic performance of low pressure axial fan rotors with different blade chord length and radial load distribution

NASA Astrophysics Data System (ADS)

Carolus, Thomas

The paper examines the acoustic and aerodynamic performance of low-pressure axial fan rotors with a hub/tip ratio of 0.45. Six rotors were designed for the same working point by means of the well-known airfoil theory. The condition of an equilibrium between the static pressure gradient and the centrifugal forces is maintained. All rotors have unequally spaced blades to diminish tonal noise. The rotors are tested in a short cylindrical housing without guide vanes. All rotors show very similar flux-pressure difference characteristics. The peak efficiency and the noise performance is considerably influenced by the chosen blade design. The aerodynamically and acoustically optimal rotor is the one with the reduced load at the hub and increased load in the tip region under satisfied equilibrium conditions. It runs at the highest aerodynamic efficiency, and its noise spectrum is fairly smooth. The overall sound pressure level of this rotor is up to 8 dB (A) lower compared to the other rotors under consideration.

The steady aerodynamics of aerofoils with porosity gradients.

PubMed

Hajian, Rozhin; Jaworski, Justin W

2017-09-01

This theoretical study determines the aerodynamic loads on an aerofoil with a prescribed porosity distribution in a steady incompressible flow. A Darcy porosity condition on the aerofoil surface furnishes a Fredholm integral equation for the pressure distribution, which is solved exactly and generally as a Riemann-Hilbert problem provided that the porosity distribution is Hölder-continuous. The Hölder condition includes as a subset any continuously differentiable porosity distributions that may be of practical interest. This formal restriction on the analysis is examined by a class of differentiable porosity distributions that approach a piecewise, discontinuous function in a certain parametric limit. The Hölder-continuous solution is verified in this limit against analytical results for partially porous aerofoils in the literature. Finally, a comparison made between the new theoretical predictions and experimental measurements of SD7003 aerofoils presented in the literature. Results from this analysis may be integrated into a theoretical framework to optimize turbulence noise suppression with minimal impact to aerodynamic performance.

Sunspots and the physics of magnetic flux tubes. III - Aerodynamic lift

NASA Technical Reports Server (NTRS)

Parker, E. N.

1979-01-01

The aerodynamic lift exerted on a magnetic flux tube by the asymmetric flow around the two sides of the tube is calculated as part of an investigation of the physics of solar flux tubes. The general hydrodynamic forces on a rigid circular cylinder in a nonuniform flow of an ideal fluid are derived from the first derivatives of the velocity field. Aerodynamic lift in a radial nonuniform flow is found to act in the direction of the flow, toward the region of increased flow velocity, while in a shear flow, lift is perpendicular to the free stream and directed toward increasing flow velocity. For a general, three dimensional, large-scale stationary incompressible equilibrium flow, an expression is also derived relating the lift per unit length to the dynamical pressure, cylinder radius and the gradient of the free-stream velocity. Evidence from an asymmetric airfoil in a uniform flow indicates that lift is enhanced in a real fluid in the presence of turbulence.

Viscous Aerodynamic Shape Optimization with Installed Propulsion Effects

NASA Technical Reports Server (NTRS)

Heath, Christopher M.; Seidel, Jonathan A.; Rallabhandi, Sriram K.

2017-01-01

Aerodynamic shape optimization is demonstrated to tailor the under-track pressure signature of a conceptual low-boom supersonic aircraft. Primarily, the optimization reduces nearfield pressure waveforms induced by propulsion integration effects. For computational efficiency, gradient-based optimization is used and coupled to the discrete adjoint formulation of the Reynolds-averaged Navier Stokes equations. The engine outer nacelle, nozzle, and vertical tail fairing are axi-symmetrically parameterized, while the horizontal tail is shaped using a wing-based parameterization. Overall, 48 design variables are coupled to the geometry and used to deform the outer mold line. During the design process, an inequality drag constraint is enforced to avoid major compromise in aerodynamic performance. Linear elastic mesh morphing is used to deform volume grids between design iterations. The optimization is performed at Mach 1.6 cruise, assuming standard day altitude conditions at 51,707-ft. To reduce uncertainty, a coupled thermodynamic engine cycle model is employed that captures installed inlet performance effects on engine operation.

The steady aerodynamics of aerofoils with porosity gradients

NASA Astrophysics Data System (ADS)

Hajian, Rozhin; Jaworski, Justin W.

2017-09-01

This theoretical study determines the aerodynamic loads on an aerofoil with a prescribed porosity distribution in a steady incompressible flow. A Darcy porosity condition on the aerofoil surface furnishes a Fredholm integral equation for the pressure distribution, which is solved exactly and generally as a Riemann-Hilbert problem provided that the porosity distribution is Hölder-continuous. The Hölder condition includes as a subset any continuously differentiable porosity distributions that may be of practical interest. This formal restriction on the analysis is examined by a class of differentiable porosity distributions that approach a piecewise, discontinuous function in a certain parametric limit. The Hölder-continuous solution is verified in this limit against analytical results for partially porous aerofoils in the literature. Finally, a comparison made between the new theoretical predictions and experimental measurements of SD7003 aerofoils presented in the literature. Results from this analysis may be integrated into a theoretical framework to optimize turbulence noise suppression with minimal impact to aerodynamic performance.

A program to compute three-dimensional subsonic unsteady aerodynamic characteristics using the doublet lattice method, L216 (DUBFLEX). Volume 2: Supplemental system design and maintenance document

NASA Technical Reports Server (NTRS)

Harrison, B. A.; Richard, M.

1979-01-01

The information necessary for execution of the digital computer program L216 on the CDC 6600 is described. L216 characteristics are based on the doublet lattice method. Arbitrary aerodynamic configurations may be represented with combinations of nonplanar lifting surfaces composed of finite constant pressure panel elements, and axially summetric slender bodies composed of constant pressure line elements. Program input consists of configuration geometry, aerodynamic parameters, and modal data; output includes element geometry, pressure difference distributions, integrated aerodynamic coefficients, stability derivatives, generalized aerodynamic forces, and aerodynamic influence coefficient matrices. Optionally, modal data may be input on magnetic field (tape or disk), and certain geometric and aerodynamic output may be saved for subsequent use.

Novel Aerodynamic Design for Formula SAE Vehicles

NASA Astrophysics Data System (ADS)

Sentongo, Samuel; Carter, Austin; Cecil, Christopher; Feier, Ioan

2017-11-01

This paper identifies and evaluates the design characteristics of a novel airfoil that harnesses the Magnus Effect, applying a moving-surface boundary-layer control (MSBC) method to a Formula SAE Vehicle. The MSBC minimizes adverse pressure gradient and delays boundary layer separation through the use of a conveyor belt that interacts with the airfoil boundary layer. The MSBC allows dynamic control of the aerodynamic coefficients by variation of the belt speed, minimizing drag in high speed straights and maximizing downforce during vehicle cornering. A conveyer belt wing measuring approximately 0.9 x 0.9m in planform was designed and built to test the mechanical setup for such a MSBC wing. This study follows the relationship between inputted power and outputted surface velocity, with the goal being to maximize speed output vs. power input. The greatest hindrance to maximizing speed output is friction among belts, rollers, and stationary members. The maximum belt speed achieved during testing was 5.9 m/s with a power input of 48.8 W, which corresponds to 45.8 N of downforce based on 2D CFD results. Ongoing progress on this project is presented. United States Air Force Academy.

Application of the adjoint optimisation of shock control bump for ONERA-M6 wing

NASA Astrophysics Data System (ADS)

Nejati, A.; Mazaheri, K.

2017-11-01

This article is devoted to the numerical investigation of the shock wave/boundary layer interaction (SWBLI) as the main factor influencing the aerodynamic performance of transonic bumped airfoils and wings. The numerical analysis is conducted for the ONERA-M6 wing through a shock control bump (SCB) shape optimisation process using the adjoint optimisation method. SWBLI is analyzed for both clean and bumped airfoils and wings, and it is shown how the modified wave structure originating from upstream of the SCB reduces the wave drag, by improving the boundary layer velocity profile downstream of the shock wave. The numerical simulation of the turbulent viscous flow and a gradient-based adjoint algorithm are used to find the optimum location and shape of the SCB for the ONERA-M6 airfoil and wing. Two different geometrical models are introduced for the 3D SCB, one with linear variations, and another with periodic variations. Both configurations result in drag reduction and improvement in the aerodynamic efficiency, but the periodic model is more effective. Although the three-dimensional flow structure involves much more complexities, the overall results are shown to be similar to the two-dimensional case.

System Identification and POD Method Applied to Unsteady Aerodynamics

NASA Technical Reports Server (NTRS)

Tang, Deman; Kholodar, Denis; Juang, Jer-Nan; Dowell, Earl H.

2001-01-01

The representation of unsteady aerodynamic flow fields in terms of global aerodynamic modes has proven to be a useful method for reducing the size of the aerodynamic model over those representations that use local variables at discrete grid points in the flow field. Eigenmodes and Proper Orthogonal Decomposition (POD) modes have been used for this purpose with good effect. This suggests that system identification models may also be used to represent the aerodynamic flow field. Implicit in the use of a systems identification technique is the notion that a relative small state space model can be useful in describing a dynamical system. The POD model is first used to show that indeed a reduced order model can be obtained from a much larger numerical aerodynamical model (the vortex lattice method is used for illustrative purposes) and the results from the POD and the system identification methods are then compared. For the example considered, the two methods are shown to give comparable results in terms of accuracy and reduced model size. The advantages and limitations of each approach are briefly discussed. Both appear promising and complementary in their characteristics.

Superparamagnetic iron oxide nanoparticles (SPIONs)-loaded Trojan microparticles for targeted aerosol delivery to the lung.

PubMed

Tewes, Frederic; Ehrhardt, Carsten; Healy, Anne Marie

2014-01-01

Targeted aerosol delivery to specific regions of the lung may improve therapeutic efficiency and minimise unwanted side effects. Targeted delivery could potentially be achieved with porous microparticles loaded with superparamagnetic iron oxide nanoparticles (SPIONs)-in combination with a target-directed magnetic gradient field. The aim of this study was to formulate and evaluate the aerodynamic properties of SPIONs-loaded Trojan microparticles after delivery from a dry powder inhaler. Microparticles made of SPIONs, PEG and hydroxypropyl-β-cyclodextrin (HPβCD) were formulated by spray drying and characterised by various physicochemical methods. Aerodynamic properties were evaluated using a next generation cascade impactor (NGI), with or without a magnet positioned at stage 2. Mixing appropriate proportions of SPIONs, PEG and HPβCD allowed Trojan microparticle to be formulated. These particles had a median geometric diameter of 2.8±0.3μm and were shown to be sensitive to the magnetic field induced by a magnet having a maximum energy product of 413.8kJ/m(3). However, these particles, characterised by a mass median aerodynamic diameter (MMAD) of 10.2±2.0μm, were considered to be not inhalable. The poor aerodynamic properties resulted from aggregation of the particles. The addition of (NH4)2CO3 and magnesium stearate (MgST) to the formulation improved the aerodynamic properties of the Trojan particles and resulted in a MMAD of 2.2±0.8μm. In the presence of a magnetic field on stage 2 of the NGI, the amount of particles deposited at this stage increased 4-fold from 4.8±0.7% to 19.5±3.3%. These Trojan particles appeared highly sensitive to the magnetic field and their deposition on most of the stages of the NGI was changed in the presence compared to the absence of the magnet. If loaded with a pharmaceutical active ingredient, these particles may be useful for treating localised lung disease such as cancer nodules or bacterial infectious foci. Copyright © 2013 Elsevier B.V. All rights reserved.

Parametric Deformation of Discrete Geometry for Aerodynamic Shape Design

NASA Technical Reports Server (NTRS)

Anderson, George R.; Aftosmis, Michael J.; Nemec, Marian

2012-01-01

We present a versatile discrete geometry manipulation platform for aerospace vehicle shape optimization. The platform is based on the geometry kernel of an open-source modeling tool called Blender and offers access to four parametric deformation techniques: lattice, cage-based, skeletal, and direct manipulation. Custom deformation methods are implemented as plugins, and the kernel is controlled through a scripting interface. Surface sensitivities are provided to support gradient-based optimization. The platform architecture allows the use of geometry pipelines, where multiple modelers are used in sequence, enabling manipulation difficult or impossible to achieve with a constructive modeler or deformer alone. We implement an intuitive custom deformation method in which a set of surface points serve as the design variables and user-specified constraints are intrinsically satisfied. We test our geometry platform on several design examples using an aerodynamic design framework based on Cartesian grids. We examine inverse airfoil design and shape matching and perform lift-constrained drag minimization on an airfoil with thickness constraints. A transport wing-fuselage integration problem demonstrates the approach in 3D. In a final example, our platform is pipelined with a constructive modeler to parabolically sweep a wingtip while applying a 1-G loading deformation across the wingspan. This work is an important first step towards the larger goal of leveraging the investment of the graphics industry to improve the state-of-the-art in aerospace geometry tools.

An initial investigation into methods of computing transonic aerodynamic sensitivity coefficients

NASA Technical Reports Server (NTRS)

Carlson, Leland A.

1992-01-01

Research conducted during the period from July 1991 through December 1992 is covered. A method based upon the quasi-analytical approach was developed for computing the aerodynamic sensitivity coefficients of three dimensional wings in transonic and subsonic flow. In addition, the method computes for comparison purposes the aerodynamic sensitivity coefficients using the finite difference approach. The accuracy and validity of the methods are currently under investigation.

A program to compute three-dimensional subsonic unsteady aerodynamic characteristics using the doublet lattic method, L216 (DUBFLX). Volume 1: Engineering and usage

NASA Technical Reports Server (NTRS)

Richard, M.; Harrison, B. A.

1979-01-01

The program input presented consists of configuration geometry, aerodynamic parameters, and modal data; output includes element geometry, pressure difference distributions, integrated aerodynamic coefficients, stability derivatives, generalized aerodynamic forces, and aerodynamic influence coefficient matrices. Optionally, modal data may be input on magnetic file (tape or disk), and certain geometric and aerodynamic output may be saved for subsequent use.

NASA aerodynamics program

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.

Comparison of a discrete steepest ascent method with the continuous steepest ascent method for optimal programing

NASA Technical Reports Server (NTRS)

Childs, A. G.

1971-01-01

A discrete steepest ascent method which allows controls which are not piecewise constant (for example, it allows all continuous piecewise linear controls) was derived for the solution of optimal programming problems. This method is based on the continuous steepest ascent method of Bryson and Denham and new concepts introduced by Kelley and Denham in their development of compatible adjoints for taking into account the effects of numerical integration. The method is a generalization of the algorithm suggested by Canon, Cullum, and Polak with the details of the gradient computation given. The discrete method was compared with the continuous method for an aerodynamics problem for which an analytic solution is given by Pontryagin's maximum principle, and numerical results are presented. The discrete method converges more rapidly than the continuous method at first, but then for some undetermined reason, loses its exponential convergence rate. A comparsion was also made for the algorithm of Canon, Cullum, and Polak using piecewise constant controls. This algorithm is very competitive with the continuous algorithm.

Adaptive momentum management for large space structures

NASA Technical Reports Server (NTRS)

Hahn, E.

1987-01-01

Momentum management is discussed for a Large Space Structure (LSS) with the structure selected configuration being the Initial Orbital Configuration (IOC) of the dual keel space station. The external forces considered were gravity gradient and aerodynamic torques. The goal of the momentum management scheme developed is to remove the bias components of the external torques and center the cyclic components of the stored angular momentum. The scheme investigated is adaptive to uncertainties of the inertia tensor and requires only approximate knowledge of principle moments of inertia. Computational requirements are minimal and should present no implementation problem in a flight type computer and the method proposed is shown to be effective in the presence of attitude control bandwidths as low as .01 radian/sec.

Adaptive momentum management for the dual keel Space Station

NASA Technical Reports Server (NTRS)

Hopkins, M.; Hahn, E.

1987-01-01

The report discusses momentum management for a large space structure with the structure selected configuration being the Initial Orbital Configuration of the dual-keel Space Station. The external torques considered were gravity gradient and aerodynamic torques. The goal of the momentum management scheme developed is to remove the bias components of the external torques and center the cyclic components of the stored angular momentum. The scheme investigated is adaptive to uncertainties of the inertia tensor and requires only approximate knowledge of principal moments of inertia. Computational requirements are minimal and should present no implementation problem in a flight-type computer. The method proposed is shown to be effective in the presence of attitude control bandwidths as low as 0.01 radian/sec.

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 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.

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.

RANS simulations of wind turbine wakes: optimal tuning of turbulence closure and aerodynamic loads from LiDAR and SCADA data

NASA Astrophysics Data System (ADS)

Letizia, Stefano; Puccioni, Matteo; Zhan, Lu; Viola, Francesco; Camarri, Simone; Iungo, Giacomo Valerio

2017-11-01

Numerical simulations of wakes produced by utility-scale wind turbines still present challenges related to the variability of the atmospheric conditions and, in the most of the cases, the lack of information about the geometry and aerodynamic performance of the wind turbine blades. In order to overcome the mentioned difficulties, we propose a RANS solver for which turbine aerodynamic forcing and turbulence closure are calibrated through LiDAR and SCADA data acquired for an onshore wind farm. The wind farm under examination is located in North Texas over a relatively flat terrain. The experimental data are leveraged to maximize accuracy of the RANS predictions in terms of wake velocity field and power capture for different atmospheric stability conditions and settings of the wind turbines. The optimization of the RANS parameters is performed through an adjoint-RANS formulation and a gradient-based procedure. The optimally-tuned aerodynamic forcing and turbulence closure are then analyzed in order to investigate effects of the atmospheric stability on the evolution of wind turbine wakes and power performance. The proposed RANS solver has low computational costs comparable to those of wake engineering models, which make it a compelling tool for wind farm control and optimization. Acknowledgments: NSF I/UCRC WindSTAR IIP 1362033 and TACC.

Application of an online ion-chromatography-based instrument for gradient flux measurements of speciated nitrogen and sulfur

NASA Astrophysics Data System (ADS)

Rumsey, Ian C.; Walker, John T.

2016-06-01

The dry component of total nitrogen and sulfur atmospheric deposition remains uncertain. The lack of measurements of sufficient chemical speciation and temporal extent make it difficult to develop accurate mass budgets and sufficient process level detail is not available to improve current air-surface exchange models. Over the past decade, significant advances have been made in the development of continuous air sampling measurement techniques, resulting with instruments of sufficient sensitivity and temporal resolution to directly quantify air-surface exchange of nitrogen and sulfur compounds. However, their applicability is generally restricted to only one or a few of the compounds within the deposition budget. Here, the performance of the Monitor for AeRosols and GAses in ambient air (MARGA 2S), a commercially available online ion-chromatography-based analyzer is characterized for the first time as applied for air-surface exchange measurements of HNO3, NH3, NH4+, NO3-, SO2 and SO42-. Analytical accuracy and precision are assessed under field conditions. Chemical concentrations gradient precision are determined at the same sampling site. Flux uncertainty measured by the aerodynamic gradient method is determined for a representative 3-week period in fall 2012 over a grass field. Analytical precision and chemical concentration gradient precision were found to compare favorably in comparison to previous studies. During the 3-week period, percentages of hourly chemical concentration gradients greater than the corresponding chemical concentration gradient detection limit were 86, 42, 82, 73, 74 and 69 % for NH3, NH4+, HNO3, NO3-, SO2 and SO42-, respectively. As expected, percentages were lowest for aerosol species, owing to their relatively low deposition velocities and correspondingly smaller gradients relative to gas phase species. Relative hourly median flux uncertainties were 31, 121, 42, 43, 67 and 56 % for NH3, NH4+, HNO3, NO3-, SO2 and SO42-, respectively. Flux uncertainty is dominated by uncertainty in the chemical concentrations gradients during the day but uncertainty in the chemical concentration gradients and transfer velocity are of the same order at night. Results show the instrument is sufficiently precise for flux gradient applications.

Theoretical and experimental investigation of supersonic aerodynamic characteristics of a twin-fuselage concept

NASA Technical Reports Server (NTRS)

Wood, R. M.; Miller, D. S.; Brentner, K. S.

1983-01-01

A theoretical and experimental investigation has been conducted to evaluate the fundamental supersonic aerodynamic characteristics of a generic twin-body model at a Mach number of 2.70. Results show that existing aerodynamic prediction methods are adequate for making preliminary aerodynamic estimates.

Aerodynamic analysis of Pegasus - Computations vs reality

NASA Technical Reports Server (NTRS)

Mendenhall, Michael R.; Lesieutre, Daniel J.; Whittaker, C. H.; Curry, Robert E.; Moulton, Bryan

1993-01-01

Pegasus, a three-stage, air-launched, winged space booster was developed to provide fast and efficient commercial launch services for small satellites. The aerodynamic design and analysis of Pegasus was conducted without benefit of wind tunnel tests using only computational aerodynamic and fluid dynamic methods. Flight test data from the first two operational flights of Pegasus are now available, and they provide an opportunity to validate the accuracy of the predicted pre-flight aerodynamic characteristics. Comparisons of measured and predicted flight characteristics are presented and discussed. Results show that the computational methods provide reasonable aerodynamic design information with acceptable margins. Post-flight analyses illustrate certain areas in which improvements are desired.

Aerodynamic design using numerical optimization

NASA Technical Reports Server (NTRS)

Murman, E. M.; Chapman, G. T.

1983-01-01

The procedure of using numerical optimization methods coupled with computational fluid dynamic (CFD) codes for the development of an aerodynamic design is examined. Several approaches that replace wind tunnel tests, develop pressure distributions and derive designs, or fulfill preset design criteria are presented. The method of Aerodynamic Design by Numerical Optimization (ADNO) is described and illustrated with examples.

Calculation of subsonic and supersonic steady and unsteady aerodynamic forces using velocity potential aerodynamic elements

NASA Technical Reports Server (NTRS)

Haviland, J. K.; Yoo, Y. S.

1976-01-01

Expressions for calculation of subsonic and supersonic, steady and unsteady aerodynamic forces are derived, using the concept of aerodynamic elements applied to the downwash velocity potential method. Aerodynamic elements can be of arbitrary out of plane polygon shape, although numerical calculations are restricted to rectangular elements, and to the steady state case in the supersonic examples. It is suggested that the use of conforming, in place of rectangular elements, would give better results. Agreement with results for subsonic oscillating T tails is fair, but results do not converge as the number of collocation points is increased. This appears to be due to the form of expression used in the calculations. The methods derived are expected to facilitate automated flutter analysis on the computer. In particular, the aerodynamic element concept is consistent with finite element methods already used for structural analysis. The method is universal for the complete Mach number range, and, finally, the calculations can be arranged so that they do not have to be repeated completely for every reduced frequency.

Nonlinear Unsteady Aerodynamic Modeling Using Wind Tunnel and Computational Data

NASA Technical Reports Server (NTRS)

Murphy, Patrick C.; Klein, Vladislav; Frink, Neal T.

2016-01-01

Extensions to conventional aircraft aerodynamic models are required to adequately predict responses when nonlinear unsteady flight regimes are encountered, especially at high incidence angles and under maneuvering conditions. For a number of reasons, such as loss of control, both military and civilian aircraft may extend beyond normal and benign aerodynamic flight conditions. In addition, military applications may require controlled flight beyond the normal envelope, and civilian flight may require adequate recovery or prevention methods from these adverse conditions. These requirements have led to the development of more general aerodynamic modeling methods and provided impetus for researchers to improve both techniques and the degree of collaboration between analytical and experimental research efforts. In addition to more general mathematical model structures, dynamic test methods have been designed to provide sufficient information to allow model identification. This paper summarizes research to develop a modeling methodology appropriate for modeling aircraft aerodynamics that include nonlinear unsteady behaviors using both experimental and computational test methods. This work was done at Langley Research Center, primarily under the NASA Aviation Safety Program, to address aircraft loss of control, prevention, and recovery aerodynamics.

NASA aerodynamics program

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.

Rarefaction Effects in Low Reynolds Number Subsonic and Transonic Aerodynamics

NASA Astrophysics Data System (ADS)

Pekardan, Cem

The quantification of rarefaction effects for low Reynolds number (Re<10,000) transonic (M=0.8) flows is essential for the aerodynamic design of vehicles moving in vacuum environments approaching slip regime. Potential future applications in these conditions include low-pressure high-speed ground transportation, high-altitude unmanned aerial vehicles, Martian aircraft and rotorcraft. For the quantification of rarefaction effects, the NACA 0012 airfoil was analyzed using the traditional Navier-Stokes equations in the low-Reynolds transonic regime. The results were compared to the deterministic solution of the ES-BGK type Boltzmann equation with the Runge-Kutta Discontinuous Galerkin Method (RKDG). Numerical simulations using these computational methods were compared to the electron beam fluorescence experiments at a Re=73 and a M=0.8, and it was observed that the numerical solution of the ES-BGK model using the RKDG method with 3rd order accuracy is computationally the most efficient. It was also shown that when the Reynolds number of the flow decreased from 10,000 to 1,000, slip effects become dominant. The flow becomes fully rarefied at Re=10. Furthermore, rarefaction effects were quantified for the NACA 0007 and the NACA 2407 at 0 and 10 degrees of angle of attack to investigate the effects of thickness, camber, and the angle of attack. It was observed that flow separation due to increase in thickness resulted in higher rarefaction effects. It was concluded that thin airfoils with very smooth shape changes minimize continuum breakdown / rarefaction effects. Rarefied gas phenomena that only appear in low pressures (such as thermal effects) can be exploited for performance enhancement of applications in slightly rarefied aerodynamics. In this study, feasibility and advantages of using thermal control to reduce drag and mitigate vortex shedding for airfoils are studied. NACA 0012 airfoil with a temperature difference applied between the upper and the lower surface is simulated in the continuum regime with a Navier-Stokes solver and compared to experimental data for verification of parameters and turbulence modelling. At lower pressures, an elevated temperature on the bottom surface of the airfoil is investigated to create lift and understand the rarefaction effects. Continuum NS results were compared to the rarefied ES-BGK solver for the rarefaction effects. It was shown that an elevated temperature enhances the lift by 25 % and reduces the drag at high angles of attack. In the second part, a temperature gradient on the upper surface is applied and it was seen that drag is reduced by 4 % and vortex shedding frequency is reduced due to gradients introduced in the flow by thermal transpiration.

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 percent 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 aerodynamic 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 -0.14 percent.

Airfoil Design and Optimization by the One-Shot Method

NASA Technical Reports Server (NTRS)

Kuruvila, G.; Taasan, Shlomo; Salas, M. D.

1995-01-01

An efficient numerical approach for the design of optimal aerodynamic shapes is presented in this paper. The objective of any optimization problem is to find the optimum of a cost function subject to a certain state equation (governing equation of the flow field) and certain side constraints. As in classical optimal control methods, the present approach introduces a costate variable (Lagrange multiplier) to evaluate the gradient of the cost function. High efficiency in reaching the optimum solution is achieved by using a multigrid technique and updating the shape in a hierarchical manner such that smooth (low-frequency) changes are done separately from high-frequency changes. Thus, the design variables are changed on a grid where their changes produce nonsmooth (high-frequency) perturbations that can be damped efficiently by the multigrid. The cost of solving the optimization problem is approximately two to three times the cost of the equivalent analysis problem.

Airfoil optimization by the one-shot method

NASA Technical Reports Server (NTRS)

Kuruvila, G.; Taasan, Shlomo; Salas, M. D.

1994-01-01

An efficient numerical approach for the design of optimal aerodynamic shapes is presented in this paper. The objective of any optimization problem is to find the optimum of a cost function subject to a certain state equation (Governing equation of the flow field) and certain side constraints. As in classical optimal control methods, the present approach introduces a costate variable (Language multiplier) to evaluate the gradient of the cost function. High efficiency in reaching the optimum solution is achieved by using a multigrid technique and updating the shape in a hierarchical manner such that smooth (low-frequency) changes are done separately from high-frequency changes. Thus, the design variables are changed on a grid where their changes produce nonsmooth (high-frequency) perturbations that can be damped efficiently by the multigrid. The cost of solving the optimization problem is approximately two to three times the cost of the equivalent analysis problem.

An approach to constrained aerodynamic design with application to airfoils

NASA Technical Reports Server (NTRS)

Campbell, Richard L.

1992-01-01

An approach was developed for incorporating flow and geometric constraints into the Direct Iterative Surface Curvature (DISC) design method. In this approach, an initial target pressure distribution is developed using a set of control points. The chordwise locations and pressure levels of these points are initially estimated either from empirical relationships and observed characteristics of pressure distributions for a given class of airfoils or by fitting the points to an existing pressure distribution. These values are then automatically adjusted during the design process to satisfy the flow and geometric constraints. The flow constraints currently available are lift, wave drag, pitching moment, pressure gradient, and local pressure levels. The geometric constraint options include maximum thickness, local thickness, leading-edge radius, and a 'glove' constraint involving inner and outer bounding surfaces. This design method was also extended to include the successive constraint release (SCR) approach to constrained minimization.

Domain decomposition for aerodynamic and aeroacoustic analyses, and optimization

NASA Technical Reports Server (NTRS)

Baysal, Oktay

1995-01-01

The overarching theme was the domain decomposition, which intended to improve the numerical solution technique for the partial differential equations at hand; in the present study, those that governed either the fluid flow, or the aeroacoustic wave propagation, or the sensitivity analysis for a gradient-based optimization. The role of the domain decomposition extended beyond the original impetus of discretizing geometrical complex regions or writing modular software for distributed-hardware computers. It induced function-space decompositions and operator decompositions that offered the valuable property of near independence of operator evaluation tasks. The objectives have gravitated about the extensions and implementations of either the previously developed or concurrently being developed methodologies: (1) aerodynamic sensitivity analysis with domain decomposition (SADD); (2) computational aeroacoustics of cavities; and (3) dynamic, multibody computational fluid dynamics using unstructured meshes.

Enveloping Aerodynamic Decelerator

NASA Technical Reports Server (NTRS)

Nock, Kerry T. (Inventor); Aaron, Kim M. (Inventor); McRonald, Angus D. (Inventor); Gates, Kristin L. (Inventor)

2018-01-01

An inflatable aerodynamic deceleration method and system is provided for use with an atmospheric entry payload. The inflatable aerodynamic decelerator includes an inflatable envelope and an inflatant, wherein the inflatant is configured to fill the inflatable envelope to an inflated state such that the inflatable envelope surrounds the atmospheric entry payload, causing aerodynamic forces to decelerate the atmospheric entry payload.

Observations Regarding Use of Advanced CFD Analysis, Sensitivity Analysis, and Design Codes in MDO

NASA Technical Reports Server (NTRS)

Newman, Perry A.; Hou, Gene J. W.; Taylor, Arthur C., III

1996-01-01

Observations regarding the use of advanced computational fluid dynamics (CFD) analysis, sensitivity analysis (SA), and design codes in gradient-based multidisciplinary design optimization (MDO) reflect our perception of the interactions required of CFD and our experience in recent aerodynamic design optimization studies using CFD. Sample results from these latter studies are summarized for conventional optimization (analysis - SA codes) and simultaneous analysis and design optimization (design code) using both Euler and Navier-Stokes flow approximations. The amount of computational resources required for aerodynamic design using CFD via analysis - SA codes is greater than that required for design codes. Thus, an MDO formulation that utilizes the more efficient design codes where possible is desired. However, in the aerovehicle MDO problem, the various disciplines that are involved have different design points in the flight envelope; therefore, CFD analysis - SA codes are required at the aerodynamic 'off design' points. The suggested MDO formulation is a hybrid multilevel optimization procedure that consists of both multipoint CFD analysis - SA codes and multipoint CFD design codes that perform suboptimizations.

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.

Recent NASA Research on Aerodynamic Modeling of Post-Stall and Spin Dynamics of Large Transport Airplanes

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.

Efficient Gradient-Based Shape Optimization Methodology Using Inviscid/Viscous CFD

NASA Technical Reports Server (NTRS)

Baysal, Oktay

1997-01-01

The formerly developed preconditioned-biconjugate-gradient (PBCG) solvers for the analysis and the sensitivity equations had resulted in very large error reductions per iteration; quadratic convergence was achieved whenever the solution entered the domain of attraction to the root. Its memory requirement was also lower as compared to a direct inversion solver. However, this memory requirement was high enough to preclude the realistic, high grid-density design of a practical 3D geometry. This limitation served as the impetus to the first-year activity (March 9, 1995 to March 8, 1996). Therefore, the major activity for this period was the development of the low-memory methodology for the discrete-sensitivity-based shape optimization. This was accomplished by solving all the resulting sets of equations using an alternating-direction-implicit (ADI) approach. The results indicated that shape optimization problems which required large numbers of grid points could be resolved with a gradient-based approach. Therefore, to better utilize the computational resources, it was recommended that a number of coarse grid cases, using the PBCG method, should initially be conducted to better define the optimization problem and the design space, and obtain an improved initial shape. Subsequently, a fine grid shape optimization, which necessitates using the ADI method, should be conducted to accurately obtain the final optimized shape. The other activity during this period was the interaction with the members of the Aerodynamic and Aeroacoustic Methods Branch of Langley Research Center during one stage of their investigation to develop an adjoint-variable sensitivity method using the viscous flow equations. This method had algorithmic similarities to the variational sensitivity methods and the control-theory approach. However, unlike the prior studies, it was considered for the three-dimensional, viscous flow equations. The major accomplishment in the second period of this project (March 9, 1996 to March 8, 1997) was the extension of the shape optimization methodology for the Thin-Layer Navier-Stokes equations. Both the Euler-based and the TLNS-based analyses compared with the analyses obtained using the CFL3D code. The sensitivities, again from both levels of the flow equations, also compared very well with the finite-differenced sensitivities. A fairly large set of shape optimization cases were conducted to study a number of issues previously not well understood. The testbed for these cases was the shaping of an arrow wing in Mach 2.4 flow. All the final shapes, obtained either from a coarse-grid-based or a fine-grid-based optimization, using either a Euler-based or a TLNS-based analysis, were all re-analyzed using a fine-grid, TLNS solution for their function evaluations. This allowed for a more fair comparison of their relative merits. From the aerodynamic performance standpoint, the fine-grid TLNS-based optimization produced the best shape, and the fine-grid Euler-based optimization produced the lowest cruise efficiency.

Automated Simulation Updates based on Flight Data

NASA Technical Reports Server (NTRS)

Morelli, Eugene A.; Ward, David G.

2007-01-01

A statistically-based method for using flight data to update aerodynamic data tables used in flight simulators is explained and demonstrated. A simplified wind-tunnel aerodynamic database for the F/A-18 aircraft is used as a starting point. Flight data from the NASA F-18 High Alpha Research Vehicle (HARV) is then used to update the data tables so that the resulting aerodynamic model characterizes the aerodynamics of the F-18 HARV. Prediction cases are used to show the effectiveness of the automated method, which requires no ad hoc adjustments by the analyst.

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.

A General Interface Method for Aeroelastic Analysis of Aircraft

NASA Technical Reports Server (NTRS)

Tzong, T.; Chen, H. H.; Chang, K. C.; Wu, T.; Cebeci, T.

1996-01-01

The aeroelastic analysis of an aircraft requires an accurate and efficient procedure to couple aerodynamics and structures. The procedure needs an interface method to bridge the gap between the aerodynamic and structural models in order to transform loads and displacements. Such an interface method is described in this report. This interface method transforms loads computed by any aerodynamic code to a structural finite element (FE) model and converts the displacements from the FE model to the aerodynamic model. The approach is based on FE technology in which virtual work is employed to transform the aerodynamic pressures into FE nodal forces. The displacements at the FE nodes are then converted back to aerodynamic grid points on the aircraft surface through the reciprocal theorem in structural engineering. The method allows both high and crude fidelities of both models and does not require an intermediate modeling. In addition, the method performs the conversion of loads and displacements directly between individual aerodynamic grid point and its corresponding structural finite element and, hence, is very efficient for large aircraft models. This report also describes the application of this aero-structure interface method to a simple wing and an MD-90 wing. The results show that the aeroelastic effect is very important. For the simple wing, both linear and nonlinear approaches are used. In the linear approach, the deformation of the structural model is considered small, and the loads from the deformed aerodynamic model are applied to the original geometry of the structure. In the nonlinear approach, the geometry of the structure and its stiffness matrix are updated in every iteration and the increments of loads from the previous iteration are applied to the new structural geometry in order to compute the displacement increments. Additional studies to apply the aero-structure interaction procedure to more complicated geometry will be conducted in the second phase of the present contract.

A General Simulation Method for Multiple Bodies in Proximate Flight

NASA Technical Reports Server (NTRS)

Meakin, Robert L.

2003-01-01

Methods of unsteady aerodynamic simulation for an arbitrary number of independent bodies flying in close proximity are considered. A novel method to efficiently detect collision contact points is described. A method to compute body trajectories in response to aerodynamic loads, applied loads, and inter-body collisions is also given. The physical correctness of the methods are verified by comparison to a set of analytic solutions. The methods, combined with a Navier-Stokes solver, are used to demonstrate the possibility of predicting the unsteady aerodynamics and flight trajectories of moving bodies that involve rigid-body collisions.

Multigrid Methods for Aerodynamic Problems in Complex Geometries

NASA Technical Reports Server (NTRS)

Caughey, David A.

1995-01-01

Work has been directed at the development of efficient multigrid methods for the solution of aerodynamic problems involving complex geometries, including the development of computational methods for the solution of both inviscid and viscous transonic flow problems. The emphasis is on problems of complex, three-dimensional geometry. The methods developed are based upon finite-volume approximations to both the Euler and the Reynolds-Averaged Navier-Stokes equations. The methods are developed for use on multi-block grids using diagonalized implicit multigrid methods to achieve computational efficiency. The work is focused upon aerodynamic problems involving complex geometries, including advanced engine inlets.

In vivo recording of aerodynamic force with an aerodynamic force platform: from drones to birds.

PubMed

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.

Aerodynamic optimization studies on advanced architecture computers

NASA Technical Reports Server (NTRS)

Chawla, Kalpana

1995-01-01

The approach to carrying out multi-discipline aerospace design studies in the future, especially in massively parallel computing environments, comprises of choosing (1) suitable solvers to compute solutions to equations characterizing a discipline, and (2) efficient optimization methods. In addition, for aerodynamic optimization problems, (3) smart methodologies must be selected to modify the surface shape. In this research effort, a 'direct' optimization method is implemented on the Cray C-90 to improve aerodynamic design. It is coupled with an existing implicit Navier-Stokes solver, OVERFLOW, to compute flow solutions. The optimization method is chosen such that it can accomodate multi-discipline optimization in future computations. In the work , however, only single discipline aerodynamic optimization will be included.

Thermophysical fundamentals of cyclonic recirculating heating devices

NASA Astrophysics Data System (ADS)

Karpov, S. V.; Zagoskin, A. A.

2017-10-01

This report presents the results of experimental and theoretical research of aerodynamics and convective heat transfer in cyclone devices with the new system of external recirculation of heating gas under the influence of radial pressure gradient in a heat carrier’s swirling turbulent flow. The dynamic problem of tangential velocity distribution in a clearance volume is solved at various re-circulation ratio values including limiting quantities (kr = 0; 1) and variations in cyclonic combustion chamber’s design parameters and operating conditions (Rer); the integrated calculation ratios for fundamental aerodynamic characteristics of a recirculation device are derived. The first experimental and numerical studies of convective heat transfer on internal and external surfaces of a hollow shaft in a swirling recirculation flow are derived through the instrumentality of OpenFOAM, these studies are also conducted for a setting of several cylindrical solid inserts. The external surface heat problem of a hollow cylindrical insert is solved with integral and digital methods; generalized similarity equations for the internal and external surfaces extended in range of Reynolds number are derived. The experimental data is in reasonable agreement with the derived curves and the results of mathematic modelling of convective heat transfer. Calculation recommendations for optimal selection of kr values at various ratios of their geometric characteristics and products utilization rate are obtained.

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.

Using the HARV simulation aerodynamic model to determine forebody strake aerodynamic coefficients from flight 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.

Effect of body aerodynamics on the dynamic flight stability of the hawkmoth Manduca sexta.

PubMed

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.

Application of an online ion chromatography-based instrument ...

EPA Pesticide Factsheets

In North America, the dry component of total nitrogen and sulfur deposition remains uncertain due to a lack of measurements of sufficient chemical speciation and temporal extent to develop complete annual mass budgets or of sufficient process level detail to improve current air-surface exchange models. Over the past decade, significant advances have been made in the development of continuous chemical measurement techniques with sufficient sensitivity and temporal resolution to directly quantify air-surface exchange of nitrogen and sulfur compounds. However, their applicability is generally restricted to only one or a few of the compounds within the deposition budget. We characterize for the first time the performance of the Monitor for AeRosols and GAses in ambient air (MARGA), an on-line ion chromatography-based analyzer, as applied for air-surface exchange measurements of HNO3, NH3, NH4+, NO3-, SO2 and SO42-. Analytical accuracy and precision are assessed under field conditions and total uncertainty of fluxes measured by the aerodynamic gradient method are assessed for a representative 3-week period in the fall of 2012. During this period, percentages of hourly chemical gradients larger than the corresponding gradient detection limit were 86%, 55%, 81%, 74%, 77%, and 71% for NH3, NH4+, HNO3, NO3-, SO2, and SO42-, respectively. As expected, percentages were lowest for aerosol species, owing to their relatively low deposition velocities and correspondingly sma

Aerodynamic design on high-speed trains

NASA Astrophysics Data System (ADS)

Ding, San-San; Li, Qiang; Tian, Ai-Qin; Du, Jian; Liu, Jia-Li

2016-04-01

Compared with the traditional train, the operational speed of the high-speed train has largely improved, and the dynamic environment of the train has changed from one of mechanical domination to one of aerodynamic domination. The aerodynamic problem has become the key technological challenge of high-speed trains and significantly affects the economy, environment, safety, and comfort. In this paper, the relationships among the aerodynamic design principle, aerodynamic performance indexes, and design variables are first studied, and the research methods of train aerodynamics are proposed, including numerical simulation, a reduced-scale test, and a full-scale test. Technological schemes of train aerodynamics involve the optimization design of the streamlined head and the smooth design of the body surface. Optimization design of the streamlined head includes conception design, project design, numerical simulation, and a reduced-scale test. Smooth design of the body surface is mainly used for the key parts, such as electric-current collecting system, wheel truck compartment, and windshield. The aerodynamic design method established in this paper has been successfully applied to various high-speed trains (CRH380A, CRH380AM, CRH6, CRH2G, and the Standard electric multiple unit (EMU)) that have met expected design objectives. The research results can provide an effective guideline for the aerodynamic design of high-speed trains.

Integrated multidisciplinary design optimization using discrete sensitivity analysis for geometrically complex aeroelastic configurations

NASA Astrophysics Data System (ADS)

Newman, James Charles, III

1997-10-01

The first two steps in the development of an integrated multidisciplinary design optimization procedure capable of analyzing the nonlinear fluid flow about geometrically complex aeroelastic configurations have been accomplished in the present work. For the first step, a three-dimensional unstructured grid approach to aerodynamic shape sensitivity analysis and design optimization has been developed. The advantage of unstructured grids, when compared with a structured-grid approach, is their inherent ability to discretize irregularly shaped domains with greater efficiency and less effort. Hence, this approach is ideally suited for geometrically complex configurations of practical interest. In this work the time-dependent, nonlinear Euler equations are solved using an upwind, cell-centered, finite-volume scheme. The discrete, linearized systems which result from this scheme are solved iteratively by a preconditioned conjugate-gradient-like algorithm known as GMRES for the two-dimensional cases and a Gauss-Seidel algorithm for the three-dimensional; at steady-state, similar procedures are used to solve the accompanying linear aerodynamic sensitivity equations in incremental iterative form. As shown, this particular form of the sensitivity equation makes large-scale gradient-based aerodynamic optimization possible by taking advantage of memory efficient methods to construct exact Jacobian matrix-vector products. Various surface parameterization techniques have been employed in the current study to control the shape of the design surface. Once this surface has been deformed, the interior volume of the unstructured grid is adapted by considering the mesh as a system of interconnected tension springs. Grid sensitivities are obtained by differentiating the surface parameterization and the grid adaptation algorithms with ADIFOR, an advanced automatic-differentiation software tool. To demonstrate the ability of this procedure to analyze and design complex configurations of practical interest, the sensitivity analysis and shape optimization has been performed for several two- and three-dimensional cases. In twodimensions, an initially symmetric NACA-0012 airfoil and a high-lift multielement airfoil were examined. For the three-dimensional configurations, an initially rectangular wing with uniform NACA-0012 cross-sections was optimized; in addition, a complete Boeing 747-200 aircraft was studied. Furthermore, the current study also examines the effect of inconsistency in the order of spatial accuracy between the nonlinear fluid and linear shape sensitivity equations. The second step was to develop a computationally efficient, high-fidelity, integrated static aeroelastic analysis procedure. To accomplish this, a structural analysis code was coupled with the aforementioned unstructured grid aerodynamic analysis solver. The use of an unstructured grid scheme for the aerodynamic analysis enhances the interaction compatibility with the wing structure. The structural analysis utilizes finite elements to model the wing so that accurate structural deflections may be obtained. In the current work, parameters have been introduced to control the interaction of the computational fluid dynamics and structural analyses; these control parameters permit extremely efficient static aeroelastic computations. To demonstrate and evaluate this procedure, static aeroelastic analysis results for a flexible wing in low subsonic, high subsonic (subcritical), transonic (supercritical), and supersonic flow conditions are presented.

Application of computational aerodynamics methods to the design and analysis of transport aircraft

NASA Technical Reports Server (NTRS)

Da Costa, A. L.

1978-01-01

The application and validation of several computational aerodynamic methods in the design and analysis of transport aircraft is established. An assessment is made concerning more recently developed methods that solve three-dimensional transonic flow and boundary layers on wings. Capabilities of subsonic aerodynamic methods are demonstrated by several design and analysis efforts. Among the examples cited are the B747 Space Shuttle Carrier Aircraft analysis, nacelle integration for transport aircraft, and winglet optimization. The accuracy and applicability of a new three-dimensional viscous transonic method is demonstrated by comparison of computed results to experimental data

Adjoint Algorithm for CAD-Based Shape Optimization Using a Cartesian Method

NASA Technical Reports Server (NTRS)

Nemec, Marian; Aftosmis, Michael J.

2004-01-01

Adjoint solutions of the governing flow equations are becoming increasingly important for the development of efficient analysis and optimization algorithms. A well-known use of the adjoint method is gradient-based shape optimization. Given an objective function that defines some measure of performance, such as the lift and drag functionals, its gradient is computed at a cost that is essentially independent of the number of design variables (geometric parameters that control the shape). More recently, emerging adjoint applications focus on the analysis problem, where the adjoint solution is used to drive mesh adaptation, as well as to provide estimates of functional error bounds and corrections. The attractive feature of this approach is that the mesh-adaptation procedure targets a specific functional, thereby localizing the mesh refinement and reducing computational cost. Our focus is on the development of adjoint-based optimization techniques for a Cartesian method with embedded boundaries.12 In contrast t o implementations on structured and unstructured grids, Cartesian methods decouple the surface discretization from the volume mesh. This feature makes Cartesian methods well suited for the automated analysis of complex geometry problems, and consequently a promising approach to aerodynamic optimization. Melvin et developed an adjoint formulation for the TRANAIR code, which is based on the full-potential equation with viscous corrections. More recently, Dadone and Grossman presented an adjoint formulation for the Euler equations. In both approaches, a boundary condition is introduced to approximate the effects of the evolving surface shape that results in accurate gradient computation. Central to automated shape optimization algorithms is the issue of geometry modeling and control. The need to optimize complex, "real-life" geometry provides a strong incentive for the use of parametric-CAD systems within the optimization procedure. In previous work, we presented an effective optimization framework that incorporates a direct-CAD interface. In this work, we enhance the capabilities of this framework with efficient gradient computations using the discrete adjoint method. We present details of the adjoint numerical implementation, which reuses the domain decomposition, multigrid, and time-marching schemes of the flow solver. Furthermore, we explain and demonstrate the use of CAD in conjunction with the Cartesian adjoint approach. The final paper will contain a number of complex geometry, industrially relevant examples with many design variables to demonstrate the effectiveness of the adjoint method on Cartesian meshes.

Reduction of adverse aerodynamic effects of large trucks, Volume I. Technical report

DOT National Transportation Integrated Search

1978-09-01

The overall objective of this study has been to develop methods of minimizing three aerodynamic-related phenomena: truck-induced aerodynamic disturbances, splash, and spray. An analytical methodology has been developed and used to characterize aerody...

Application of a transonic potential flow code to the static aeroelastic analysis of three-dimensional wings

NASA Technical Reports Server (NTRS)

Whitlow, W., Jr.; Bennett, R. M.

1982-01-01

Since the aerodynamic theory is nonlinear, the method requires the coupling of two iterative processes - an aerodynamic analysis and a structural analysis. A full potential analysis code, FLO22, is combined with a linear structural analysis to yield aerodynamic load distributions on and deflections of elastic wings. This method was used to analyze an aeroelastically-scaled wind tunnel model of a proposed executive-jet transport wing and an aeroelastic research wing. The results are compared with the corresponding rigid-wing analyses, and some effects of elasticity on the aerodynamic loading are noted.

The importance of aerodynamics for prosthetic limb design used by competitive cyclists with an amputation: An introduction.

PubMed

Dyer, Bryce

2015-06-01

This study introduces the importance of the aerodynamics to prosthetic limb design for athletes with either a lower-limb or upper-limb amputation. The study comprises two elements: 1) An initial experiment investigating the stability of outdoor velodrome-based field tests, and 2) An experiment evaluating the application of outdoor velodrome aerodynamic field tests to detect small-scale changes in aerodynamic drag respective of prosthetic limb componentry changes. An outdoor field-testing method is used to detect small and repeatable changes in the aerodynamic drag of an able-bodied cyclist. These changes were made at levels typical of alterations in prosthetic componentry. The field-based test method of assessment is used at a smaller level of resolution than previously reported. With a carefully applied protocol, the field test method proved to be statistically stable. The results of the field test experiments demonstrate a noticeable change in overall athlete performance. Aerodynamic refinement of artificial limbs is worthwhile for athletes looking to maximise their competitive performance. A field-testing method illustrates the importance of the aerodynamic optimisation of prosthetic limb components. The field-testing protocol undertaken in this study gives an accessible and affordable means of doing so by prosthetists and sports engineers. Using simple and accessible field-testing methods, this exploratory experiment demonstrates how small changes to riders' equipment, consummate of the scale of a small change in prosthetics componentry, can affect the performance of an athlete. Prosthetists should consider such opportunities for performance enhancement when possible. © The International Society for Prosthetics and Orthotics 2014.

A Matrix-Free Algorithm for Multidisciplinary Design Optimization

NASA Astrophysics Data System (ADS)

Lambe, Andrew Borean

Multidisciplinary design optimization (MDO) is an approach to engineering design that exploits the coupling between components or knowledge disciplines in a complex system to improve the final product. In aircraft design, MDO methods can be used to simultaneously design the outer shape of the aircraft and the internal structure, taking into account the complex interaction between the aerodynamic forces and the structural flexibility. Efficient strategies are needed to solve such design optimization problems and guarantee convergence to an optimal design. This work begins with a comprehensive review of MDO problem formulations and solution algorithms. First, a fundamental MDO problem formulation is defined from which other formulations may be obtained through simple transformations. Using these fundamental problem formulations, decomposition methods from the literature are reviewed and classified. All MDO methods are presented in a unified mathematical notation to facilitate greater understanding. In addition, a novel set of diagrams, called extended design structure matrices, are used to simultaneously visualize both data communication and process flow between the many software components of each method. For aerostructural design optimization, modern decomposition-based MDO methods cannot efficiently handle the tight coupling between the aerodynamic and structural states. This fact motivates the exploration of methods that can reduce the computational cost. A particular structure in the direct and adjoint methods for gradient computation motivates the idea of a matrix-free optimization method. A simple matrix-free optimizer is developed based on the augmented Lagrangian algorithm. This new matrix-free optimizer is tested on two structural optimization problems and one aerostructural optimization problem. The results indicate that the matrix-free optimizer is able to efficiently solve structural and multidisciplinary design problems with thousands of variables and constraints. On the aerostructural test problem formulated with thousands of constraints, the matrix-free optimizer is estimated to reduce the total computational time by up to 90% compared to conventional optimizers.

A Matrix-Free Algorithm for Multidisciplinary Design Optimization

NASA Astrophysics Data System (ADS)

Lambe, Andrew Borean

Multidisciplinary design optimization (MDO) is an approach to engineering design that exploits the coupling between components or knowledge disciplines in a complex system to improve the final product. In aircraft design, MDO methods can be used to simultaneously design the outer shape of the aircraft and the internal structure, taking into account the complex interaction between the aerodynamic forces and the structural flexibility. Efficient strategies are needed to solve such design optimization problems and guarantee convergence to an optimal design. This work begins with a comprehensive review of MDO problem formulations and solution algorithms. First, a fundamental MDO problem formulation is defined from which other formulations may be obtained through simple transformations. Using these fundamental problem formulations, decomposition methods from the literature are reviewed and classified. All MDO methods are presented in a unified mathematical notation to facilitate greater understanding. In addition, a novel set of diagrams, called extended design structure matrices, are used to simultaneously visualize both data communication and process flow between the many software components of each method. For aerostructural design optimization, modern decomposition-based MDO methods cannot efficiently handle the tight coupling between the aerodynamic and structural states. This fact motivates the exploration of methods that can reduce the computational cost. A particular structure in the direct and adjoint methods for gradient computation. motivates the idea of a matrix-free optimization method. A simple matrix-free optimizer is developed based on the augmented Lagrangian algorithm. This new matrix-free optimizer is tested on two structural optimization problems and one aerostructural optimization problem. The results indicate that the matrix-free optimizer is able to efficiently solve structural and multidisciplinary design problems with thousands of variables and constraints. On the aerostructural test problem formulated with thousands of constraints, the matrix-free optimizer is estimated to reduce the total computational time by up to 90% compared to conventional optimizers.

Aerodynamic characteristics of cruciform missiles at high angles of attack

NASA Technical Reports Server (NTRS)

Lesieutre, Daniel J.; Mendenhall, Michael R.; Nazario, Susana M.; Hemsch, Michael J.

1987-01-01

An aerodynamic prediction method for missile aerodynamic performance and preliminary design has been developed to utilize a newly available systematic fin data base and an improved equivalent angle of attack methodology. The method predicts total aerodynamic loads and individual fin forces and moments for body-tail (wing-body) and canard-body-tail configurations with cruciform fin arrangements. The data base and the prediction method are valid for angles of attack up to 45 deg, arbitrary roll angles, fin deflection angles between -40 deg and 40 deg, Mach numbers between 0.6 and 4.5, and fin aspect ratios between 0.25 and 4.0. The equivalent angle of attack concept is employed to include the effects of vorticity and geometric scaling.

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.

Free convection in the Matian atmosphere

NASA Technical Reports Server (NTRS)

Clow, G. D.; Haberle, R. M.

1990-01-01

The 'free convective' regime for the Martian atmospheric boundary layer (ABL) was investigated. This state occurs when the mean windspeed at the top of the ABL drops below some critical value U(sub c) and positive buoyant forces are present. Such forces can arise either from vertical temperature or water vapor gradients across the atmospheric surface layer. During free convection, buoyant forces drive narrow plumes that ascend to the inversion height with a return circulation consisting of broad slower-moving downdraughts. Horizontal pressure, temperature, windspeed, and water vapor fluctuations resulting form this circulation pattern can be quite large adjacent to the ground (within the surface layer). The local turbulent fluctuations cause non-zero mean surface stresses, sensible heat fluxes, and latent heat fluxes, even when the mean regional windspeed is zero. Although motions above the surface layer are insensitive to the nature of the surface, the sensible and latent heat fluxes are primarily controlled by processes within the interfacial sublayer immediately adjacent to the ground during free convection. Thus the distinction between aerodynamically smooth and rough airflow within the interfacial sublayer is more important than for the more typical situation where the mean regional windspeed is greater than U(sub c). Buoyant forces associated with water vapor gradients are particularly large on Mars at low pressures and high temperatures when the surface relative humidity is 100 percent, enhancing the likelihood of free convection under these conditions. On this basis, Ingersol postulated the evaporative heat losses from an icy surface on Mars at 237 K and current pressures would exceed the available net radiative flux at the surface, thus prohibiting ice from melting at low atmospheric pressures. Schumann has developed equations describing the horizontal fluctuations and mean vertical gradients occurring during free convection. Schumann's model was generalized to include convection driven by water vapor gradients and to include the effects of circulation above both aerodynamically smooth and rough surfaces.

Viscous-Inviscid Methods in Unsteady Aerodynamic Analysis of Bio-Inspired Morphing Wings

NASA Astrophysics Data System (ADS)

Dhruv, Akash V.

Flight has been one of the greatest realizations of human imagination, revolutionizing communication and transportation over the years. This has greatly influenced the growth of technology itself, enabling researchers to communicate and share their ideas more effectively, extending the human potential to create more sophisticated systems. While the end product of a sophisticated technology makes our lives easier, its development process presents an array of challenges in itself. In last decade, scientists and engineers have turned towards bio-inspiration to design more efficient and robust aerodynamic systems to enhance the ability of Unmanned Aerial Vehicles (UAVs) to be operated in cluttered environments, where tight maneuverability and controllability are necessary. Effective use of UAVs in domestic airspace will mark the beginning of a new age in communication and transportation. The design of such complex systems necessitates the need for faster and more effective tools to perform preliminary investigations in design, thereby streamlining the design process. This thesis explores the implementation of numerical panel methods for aerodynamic analysis of bio-inspired morphing wings. Numerical panel methods have been one of the earliest forms of computational methods for aerodynamic analysis to be developed. Although the early editions of this method performed only inviscid analysis, the algorithm has matured over the years as a result of contributions made by prominent aerodynamicists. The method discussed in this thesis is influenced by recent advancements in panel methods and incorporates both viscous and inviscid analysis of multi-flap wings. The surface calculation of aerodynamic coefficients makes this method less computationally expensive than traditional Computational Fluid Dynamics (CFD) solvers available, and thus is effective when both speed and accuracy are desired. The morphing wing design, which consists of sequential feather-like flaps installed over the upper and lower surfaces of a standard airfoil, proves to be an effective alternative to standard control surfaces by increasing the flight capability of bird-scale UAVs. The results obtained for this wing design under various flight and flap configurations provide insight into its aerodynamic behavior, which enhance the maneuverability and controllability. The overall method acts as an important tool to create an aerodynamic database to develop a distributed control system for autonomous operation of the multi-flap morphing wing, supporting the use of viscous-inviscid methods as a tool in rapid aerodynamic analysis.

Aerodynamic optimization of wind turbine rotor using CFD/AD method

NASA Astrophysics Data System (ADS)

Cao, Jiufa; Zhu, Weijun; Wang, Tongguang; Ke, Shitang

2018-05-01

The current work describes a novel technique for wind turbine rotor optimization. The aerodynamic design and optimization of wind turbine rotor can be achieved with different methods, such as the semi-empirical engineering methods and more accurate computational fluid dynamic (CFD) method. The CFD method often provides more detailed aerodynamics features during the design process. However, high computational cost limits the application, especially for rotor optimization purpose. In this paper, a CFD-based actuator disc (AD) model is used to represent turbulent flow over a wind turbine rotor. The rotor is modeled as a permeable disc of equivalent area where the forces from the blades are distributed on the circular disc. The AD model is coupled with a Reynolds Averaged Navier-Stokes (RANS) solver such that the thrust and power are simulated. The design variables are the shape parameters comprising the chord, the twist and the relative thickness of the wind turbine rotor blade. The comparative aerodynamic performance is analyzed between the original and optimized reference wind turbine rotor. The results showed that the optimization framework can be effectively and accurately utilized in enhancing the aerodynamic performance of the wind turbine rotor.

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 %.

Application of Approximate Unsteady Aerodynamics for Flutter Analysis

NASA Technical Reports Server (NTRS)

Pak, Chan-gi; Li, Wesley W.

2010-01-01

A technique for approximating the modal aerodynamic influence coefficient (AIC) matrices by using basis functions has been developed. A process for using the resulting approximated modal AIC matrix in aeroelastic analysis has also been developed. The method requires the unsteady aerodynamics in frequency domain, and this methodology can be applied to the unsteady subsonic, transonic, and supersonic aerodynamics. The flutter solution can be found by the classic methods, such as rational function approximation, k, p-k, p, root locus et cetera. The unsteady aeroelastic analysis using unsteady subsonic aerodynamic approximation is demonstrated herein. The technique presented is shown to offer consistent flutter speed prediction on an aerostructures test wing (ATW) 2 and a hybrid wing body (HWB) type of vehicle configuration with negligible loss in precision. This method computes AICs that are functions of the changing parameters being studied and are generated within minutes of CPU time instead of hours. These results may have practical application in parametric flutter analyses as well as more efficient multidisciplinary design and optimization studies.

Aerodynamic interaction between vortical wakes and lifting two-dimensional bodies

NASA Technical Reports Server (NTRS)

Stremel, Paul M.

1987-01-01

Unsteady rotor wake interactions with the empenage, tail boom, and other aerodynamic surfaces of a helicopter have a significant influence on its aerodynamic performance, the ride quality, and amount of vibration. A numerical method for computing the aerodynamic interaction between an interacting vortex wake and the viscous flow about arbitrary two-dimensional bodies has been developed to address this helicopter problem. The method solves for the flow field velocities on a body-fitted computational mesh using finite-difference techniques. The interaction of a rotor wake with the flow about a 4:1 elliptic cylinder at 45-deg incidence was calculated for a Reynolds number of 3000.

A method for the reduction of aerodynamic drag of road vehicles

NASA Technical Reports Server (NTRS)

Pamadi, Bandu N.; Taylor, Larry W.; Leary, Terrance O.

1990-01-01

A method is proposed for the reduction of the aerodynamic drag of bluff bodies, particularly for application to road transport vehicles. This technique consists of installation of panels on the forward surface of the vehicle facing the airstream. With the help of road tests, it was demonstrated that the attachment of proposed panels can reduce aerodynamic drag of road vehicles and result in significant fuel cost savings and conservation of energy resources.

Basis Function Approximation of Transonic Aerodynamic Influence Coefficient Matrix

NASA Technical Reports Server (NTRS)

Li, Wesley W.; Pak, Chan-gi

2011-01-01

A technique for approximating the modal aerodynamic influence coefficients matrices by using basis functions has been developed and validated. An application of the resulting approximated modal aerodynamic influence coefficients matrix for a flutter analysis in transonic speed regime has been demonstrated. This methodology can be applied to the unsteady subsonic, transonic, and supersonic aerodynamics. The method requires the unsteady aerodynamics in frequency-domain. The flutter solution can be found by the classic methods, such as rational function approximation, k, p-k, p, root-locus et cetera. The unsteady aeroelastic analysis for design optimization using unsteady transonic aerodynamic approximation is being demonstrated using the ZAERO flutter solver (ZONA Technology Incorporated, Scottsdale, Arizona). The technique presented has been shown to offer consistent flutter speed prediction on an aerostructures test wing 2 configuration with negligible loss in precision in transonic speed regime. These results may have practical significance in the analysis of aircraft aeroelastic calculation and could lead to a more efficient design optimization cycle.

Special Course on Engineering Methods in Aerodynamic Analysis and Design of Aircraft (Les Methodes d’Ingenierie Employee lors de l’Analyse et de la Conception des Aeronefs)

DTIC Science & Technology

1992-01-01

study assumed to the VLM-SA method (solid carve), obtained by combininglie between 15 and 75% of the local wing chord) The ex - the potential-flow...the Consultant and Lxt hange Programme and the Aeroslace Apphations Studies Programme The results oa AGARD work are reported it) the member rations...Testing AGARD CP.429, July 1988 Aerodynamics of Hypersonic, Lifting Vehicles AGARD CP-428, Novemnber 1987 Aerodynamic and Related Hydr dynamic Studies

Aerodynamics, Dimensions, Inertial Properties, and Performance of Artillery Projectiles

DTIC Science & Technology

1977-01-01

vary in ’*ese characteristics due both’o ru-mn-inv1 chancies r-ade over the years and! changes in the method of r 1a71_1faCtL rf- and of maoufacturer...Aerodynamic Coefficients All of the aerodynamic coefficients presented in this report, except for the XM712 (CLGP), were estimated by the same method and...are pre- sented in the same format. The method used is documented in Reference 22 and is available as a computer program, SPIN73, in FORTRAN. It

UHTC Research at NASA Ames

NASA Technical Reports Server (NTRS)

Johnson, Sylvia M.

2011-01-01

For enhanced aerodynamic performance. Materials for sharp leading edges can be reusable but need different properties because of geometry and very high temperatures. Require materials with significantly higher temperature capabilities, but for short duration. Current shuttle RCC leading edge materials: T approx. 1650 C. Materials for vehicles with sharp leading edges: T>2000 C. >% Figure depicts: High Temperature at Tip and Steep Temperature Gradient. Passive cooling is simplest option to manage the intense heating on sharp leading edges.

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.

Field Effects of Buoyancy on Lean Premixed Turbulent Flames

NASA Technical Reports Server (NTRS)

Cheng, R. K.; Dimalanta, R.; Wernet, M. P.; Greenberg, P. S.

2001-01-01

Buoyancy affects the entire flowfield of steady turbulent flames and this aspect of flame buoyancy coupling is largely unexplored by experiments or by theory. Open flames and flames within large confinements are free to expand and interact with the surrounding environment. In addition to fluid and combustion conditions, their aerodynamic flowfields are determined by the flame brush orientation and geometry, wake of the stabilizer, enclosure size, and of course, the gravitational field. Because the flowfield consists mainly of cold reactants (mostly in the nearfield) and hot products (mostly in the farfield), buoyancy effects are manifested in the farfield region. In upward pointing flames, an obvious effect is a favorable axial pressure gradient that accelerates the products thereby increasing the axial aerodynamic stretch rate. Intrinsic to turbulent flows, changes in mean aerodynamic stretch also couple to the fluctuating pressure field. Consequently, buoyancy can influence the turbulence intensities upstream and downstream of the flame. Flame wrinkling process, and heat release rate are also directly affected. This backward coupling mechanism is the so-called elliptic problem. To resolve the field effects of buoyancy would require the solution of three-dimensional non-linear Navier Stokes equations with full specification of the upstream, wall and downstream boundary conditions.

Investigation of Positively Curved Blade in Compressor Cascade Based on Transition Model

NASA Astrophysics Data System (ADS)

Chen, Shaowen; Lan, Yunhe; Zhou, Zhihua; Wang, Songtao

2016-06-01

Experiment and numerical simulation of flow transition in a compressor cascade with positively curved blade is carried out in a low speed. In the experimental investigation, the outlet aerodynamic parameters are measured using a five-hole aerodynamic probe, and an ink-trace flow visualization is applied to the cascade surface. The effects of transition flow on the boundary layer development, three-dimensional flow separation and aerodynamic performance are studied. The feasibility of a commercial computational fluid dynamic code is validated and the numerical results show a good agreement with experimental data. The blade-positive curving intensifies the radial force from the endwalls to the mid-span near the suction surface, which leads to the smaller scope of the intermittent region, the lesser extents of turbulence intensity and the shorter radial height of the separation bubble near the endwalls, but has little influence on the flow near the mid-span. The large passage vortex is divided into two smaller shedding vortexes under the impact of the radial pressure gradient due to the positively curved blade. The new concentrated shedding vortex results in an increase in the turbulence intensity and secondary flow loss of the corresponding region.

Calibration of a γ- Re θ transition model and its application in low-speed flows

NASA Astrophysics Data System (ADS)

Wang, YunTao; Zhang, YuLun; Meng, DeHong; Wang, GunXue; Li, Song

2014-12-01

The prediction of laminar-turbulent transition in boundary layer is very important for obtaining accurate aerodynamic characteristics with computational fluid dynamic (CFD) tools, because laminar-turbulent transition is directly related to complex flow phenomena in boundary layer and separated flow in space. Unfortunately, the transition effect isn't included in today's major CFD tools because of non-local calculations in transition modeling. In this paper, Menter's γ- Re θ transition model is calibrated and incorporated into a Reynolds-Averaged Navier-Stokes (RANS) code — Trisonic Platform (TRIP) developed in China Aerodynamic Research and Development Center (CARDC). Based on the experimental data of flat plate from the literature, the empirical correlations involved in the transition model are modified and calibrated numerically. Numerical simulation for low-speed flow of Trapezoidal Wing (Trap Wing) is performed and compared with the corresponding experimental data. It is indicated that the γ- Re θ transition model can accurately predict the location of separation-induced transition and natural transition in the flow region with moderate pressure gradient. The transition model effectively imporves the simulation accuracy of the boundary layer and aerodynamic characteristics.

Glottal Aerodynamic Measures in Women with Phonotraumatic and Nonphonotraumatic Vocal Hyperfunction

ERIC Educational Resources Information Center

Espinoza, Victor M.; Zañartu, Matías; Van Stan, Jarrad H.; Mehta, Daryush D.; Hillman, Robert E.

2017-01-01

Purpose: The purpose of this study was to determine the validity of preliminary reports showing that glottal aerodynamic measures can identify pathophysiological phonatory mechanisms for phonotraumatic and nonphonotraumatic vocal hyperfunction, which are each distinctly different from normal vocal function. Method: Glottal aerodynamic measures…

User's Guide for ECAP2D: an Euler Unsteady Aerodynamic and Aeroelastic Analysis Program for Two Dimensional Oscillating Cascades, Version 1.0

NASA Technical Reports Server (NTRS)

Reddy, T. S. R.

1995-01-01

This guide describes the input data required for using ECAP2D (Euler Cascade Aeroelastic Program-Two Dimensional). ECAP2D can be used for steady or unsteady aerodynamic and aeroelastic analysis of two dimensional cascades. Euler equations are used to obtain aerodynamic forces. The structural dynamic equations are written for a rigid typical section undergoing pitching (torsion) and plunging (bending) motion. The solution methods include harmonic oscillation method, influence coefficient method, pulse response method, and time integration method. For harmonic oscillation method, example inputs and outputs are provided for pitching motion and plunging motion. For the rest of the methods, input and output for pitching motion only are given.

Least Squares Shadowing Sensitivity Analysis of Chaotic Flow Around a Two-Dimensional Airfoil

NASA Technical Reports Server (NTRS)

Blonigan, Patrick J.; Wang, Qiqi; Nielsen, Eric J.; Diskin, Boris

2016-01-01

Gradient-based sensitivity analysis has proven to be an enabling technology for many applications, including design of aerospace vehicles. However, conventional sensitivity analysis methods break down when applied to long-time averages of chaotic systems. This breakdown is a serious limitation because many aerospace applications involve physical phenomena that exhibit chaotic dynamics, most notably high-resolution large-eddy and direct numerical simulations of turbulent aerodynamic flows. A recently proposed methodology, Least Squares Shadowing (LSS), avoids this breakdown and advances the state of the art in sensitivity analysis for chaotic flows. The first application of LSS to a chaotic flow simulated with a large-scale computational fluid dynamics solver is presented. The LSS sensitivity computed for this chaotic flow is verified and shown to be accurate, but the computational cost of the current LSS implementation is high.

Towards physiologically meaningful water-use efficiency estimates from eddy covariance data.

PubMed

Knauer, Jürgen; Zaehle, Sönke; Medlyn, Belinda E; Reichstein, Markus; Williams, Christopher A; Migliavacca, Mirco; De Kauwe, Martin G; Werner, Christiane; Keitel, Claudia; Kolari, Pasi; Limousin, Jean-Marc; Linderson, Maj-Lena

2018-02-01

Intrinsic water-use efficiency (iWUE) characterizes the physiological control on the simultaneous exchange of water and carbon dioxide in terrestrial ecosystems. Knowledge of iWUE is commonly gained from leaf-level gas exchange measurements, which are inevitably restricted in their spatial and temporal coverage. Flux measurements based on the eddy covariance (EC) technique can overcome these limitations, as they provide continuous and long-term records of carbon and water fluxes at the ecosystem scale. However, vegetation gas exchange parameters derived from EC data are subject to scale-dependent and method-specific uncertainties that compromise their ecophysiological interpretation as well as their comparability among ecosystems and across spatial scales. Here, we use estimates of canopy conductance and gross primary productivity (GPP) derived from EC data to calculate a measure of iWUE (G 1 , "stomatal slope") at the ecosystem level at six sites comprising tropical, Mediterranean, temperate, and boreal forests. We assess the following six mechanisms potentially causing discrepancies between leaf and ecosystem-level estimates of G 1 : (i) non-transpirational water fluxes; (ii) aerodynamic conductance; (iii) meteorological deviations between measurement height and canopy surface; (iv) energy balance non-closure; (v) uncertainties in net ecosystem exchange partitioning; and (vi) physiological within-canopy gradients. Our results demonstrate that an unclosed energy balance caused the largest uncertainties, in particular if it was associated with erroneous latent heat flux estimates. The effect of aerodynamic conductance on G 1 was sufficiently captured with a simple representation. G 1 was found to be less sensitive to meteorological deviations between canopy surface and measurement height and, given that data are appropriately filtered, to non-transpirational water fluxes. Uncertainties in the derived GPP and physiological within-canopy gradients and their implications for parameter estimates at leaf and ecosystem level are discussed. Our results highlight the importance of adequately considering the sources of uncertainty outlined here when EC-derived water-use efficiency is interpreted in an ecophysiological context. © 2017 John Wiley & Sons Ltd.

Derivation of aerodynamic kernel functions

NASA Technical Reports Server (NTRS)

Dowell, E. H.; Ventres, C. S.

1973-01-01

The method of Fourier transforms is used to determine the kernel function which relates the pressure on a lifting surface to the prescribed downwash within the framework of Dowell's (1971) shear flow model. This model is intended to improve upon the potential flow aerodynamic model by allowing for the aerodynamic boundary layer effects neglected in the potential flow model. For simplicity, incompressible, steady flow is considered. The proposed method is illustrated by deriving known results from potential flow theory.

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.

Recent theoretical developments and experimental studies pertinent to vortex flow aerodynamics - With a view towards design

NASA Technical Reports Server (NTRS)

Lamar, J. E.; Luckring, J. M.

1978-01-01

A review is presented of recent progress in a research program directed towards the development of an improved vortex-flow technology base. It is pointed out that separation induced vortex-flows from the leading and side edges play an important role in the high angle-of-attack aerodynamic characteristics of a wide range of modern aircraft. In the analysis and design of high-speed aircraft, a detailed knowledge of this type of separation is required, particularly with regard to critical wind loads and the stability and performance at various off-design conditions. A description of analytical methods is presented. The theoretical methods employed are divided into two classes which are dependent upon the underlying aerodynamic assumptions. One conical flow method is considered along with three different nonconical flow methods. Comparisons are conducted between the described methods and available aerodynamic data. Attention is also given to a vortex flow drag study and a vortex flow wing design using suction analogy.

Experimental and analytical research on the aerodynamics of wind driven turbines. Final report

DOE Office of Scientific and Technical Information (OSTI.GOV)

Rohrbach, C.; Wainauski, H.; Worobel, R.

1977-12-01

This aerodynamic research program was aimed at providing a reliable, comprehensive data base on a series of wind turbine models covering a broad range of the prime aerodynamic and geometric variables. Such data obtained under controlled laboratory conditions on turbines designed by the same method, of the same size, and tested in the same wind tunnel had not been available in the literature. Moreover, this research program was further aimed at providing a basis for evaluating the adequacy of existing wind turbine aerodynamic design and performance methodology, for assessing the potential of recent advanced theories and for providing a basismore » for further method development and refinement.« less

Aerodynamic laser-heated contactless furnace for neutron scattering experiments at elevated temperatures

NASA Astrophysics Data System (ADS)

Landron, Claude; Hennet, Louis; Coutures, Jean-Pierre; Jenkins, Tudor; Alétru, Chantal; Greaves, Neville; Soper, Alan; Derbyshire, Gareth

2000-04-01

Conventional radiative furnaces require sample containment that encourages contamination at elevated temperatures and generally need windows which restrict the entrance and exit solid angles required for diffraction and scattering measurements. We describe a contactless windowless furnace based on aerodynamic levitation and laser heating which has been designed for high temperature neutron scattering experiments. Data from initial experiments are reported for crystalline and amorphous oxides at temperatures up to 1900 °C, using the spallation neutron source ISIS together with our laser-heated aerodynamic levitator. Accurate reproduction of thermal expansion coefficients and radial distribution functions have been obtained, demonstrating the utility of aerodynamic levitation methods for neutron scattering methods.

Simulation of aerodynamic noise and vibration noise in hard disk drives

NASA Astrophysics Data System (ADS)

Zhu, Lei; Shen, Sheng-Nan; Li, Hui; Zhang, Guo-Qing; Cui, Fu-Hao

2018-05-01

Internal flow field characteristics of HDDs are usually influenced by the arm swing during seek operations. This, in turn, can affect aerodynamic noise and airflow-induced noise. In this paper, the dynamic mesh method is used to calculate the flow-induced vibration (FIV) by transient structure analysis and the boundary element method (BEM) is utilized to predict the vibration noise. Two operational states are considered: the arm is fixed and swinging over the disk. Both aerodynamic noise and vibration noise inside drives increase rapidly with increase in disk rotation and arm swing velocities. The largest aerodynamic noise source is always located near the arm and swung with the arm.

A comparison of the aerodynamic characteristics at transonic speeds of four wing-fuselage configurations as determined from different test techniques, 4 October 1960

NASA Technical Reports Server (NTRS)

Donlan, C. J.; Myers, B. C., II; Mattson, A. T.

1976-01-01

The high speed aerodynamic characteristics of a family of four wing-fuselage configurations of 0, 35, 45, and 60 deg sweepback were determined from transonic bump model tests that were conducted in the Langley high speed 7 by 10 foot tunnel; sting supported model tests were conducted in the Langley 8 foot high speed tunnel and in the Langley high speed 7 by 10 foot tunnel, and rocket model tests were conducted by the Langley Pilotless Aircraft Research Division. A complementary study of the effect of Mach number gradients and streamline curvature on bump results is also included. The qualitative data obtained from the various test facilities for the wing-fuselage configurations were in essential agreement as regards the relative effects of sweepback and Mach number except for drag at zero lift. Quantitatively, important differences were present.

Unsteady Computational Tests of a Non-Equilibrium

NASA Astrophysics Data System (ADS)

Jirasek, Adam; Hamlington, Peter; Lofthouse, Andrew; Usafa Collaboration; Cu Boulder Collaboration

2017-11-01

A non-equilibrium turbulence model is assessed on simulations of three practically-relevant unsteady test cases; oscillating channel flow, transonic flow around an oscillating airfoil, and transonic flow around the Benchmark Super-Critical Wing. The first case is related to piston-driven flows while the remaining cases are relevant to unsteady aerodynamics at high angles of attack and transonic speeds. Non-equilibrium turbulence effects arise in each of these cases in the form of a lag between the mean strain rate and Reynolds stresses, resulting in reduced kinetic energy production compared to classical equilibrium turbulence models that are based on the gradient transport (or Boussinesq) hypothesis. As a result of the improved representation of unsteady flow effects, the non-equilibrium model provides substantially better agreement with available experimental data than do classical equilibrium turbulence models. This suggests that the non-equilibrium model may be ideally suited for simulations of modern high-speed, high angle of attack aerodynamics problems.

Compensation of an attitude disturbance torque caused by magnetic substances in LEO satellites

NASA Astrophysics Data System (ADS)

Inamori, Takaya; Wang, Jihe; Saisutjarit, Phongsatorn; Ohsaki, Hiroyuki

This research considers an attitude disturbance torque caused by ferromagnetic substances in a LEO satellite. In most LEO satellite missions, a gravity gradient torque, solar pressure torque, aerodynamic torque, and magnetic dipole moment torque are considered for their attitude control systems, however, the effect of the ferromagnetic substances causing a disturbance torque in the geomagnetic field is not considered in previous satellite missions. The ferromagnetic substances such as iron cores of MTQs and a magnetic hysteresis damper for a passive attitude control system are used in various small satellites. These substances cause a disturbance torque which is almost the same magnitude of the dipole magnetic disturbance and the dominant disturbance in the worst cases. This research proposes a method to estimate and compensate for the effect of the ferromagnetic substances using an extended Kalman filter. From simulation results, the research concludes that the proposed method is useful and attractive for precise attitude control for LEO satellite missions.

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.

Design of control laws for flutter suppression based on the aerodynamic energy concept and comparisons with other design methods

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.

Design of control laws for flutter suppression based on the aerodynamic energy concept and comparisons with other design methods

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.

Statistical Analysis of the Uncertainty in Pre-Flight Aerodynamic Database of a Hypersonic Vehicle

NASA Astrophysics Data System (ADS)

Huh, Lynn

The objective of the present research was to develop a new method to derive the aerodynamic coefficients and the associated uncertainties for flight vehicles via post- flight inertial navigation analysis using data from the inertial measurement unit. Statistical estimates of vehicle state and aerodynamic coefficients are derived using Monte Carlo simulation. Trajectory reconstruction using the inertial navigation system (INS) is a simple and well used method. However, deriving realistic uncertainties in the reconstructed state and any associated parameters is not so straight forward. Extended Kalman filters, batch minimum variance estimation and other approaches have been used. However, these methods generally depend on assumed physical models, assumed statistical distributions (usually Gaussian) or have convergence issues for non-linear problems. The approach here assumes no physical models, is applicable to any statistical distribution, and does not have any convergence issues. The new approach obtains the statistics directly from a sufficient number of Monte Carlo samples using only the generally well known gyro and accelerometer specifications and could be applied to the systems of non-linear form and non-Gaussian distribution. When redundant data are available, the set of Monte Carlo simulations are constrained to satisfy the redundant data within the uncertainties specified for the additional data. The proposed method was applied to validate the uncertainty in the pre-flight aerodynamic database of the X-43A Hyper-X research vehicle. In addition to gyro and acceleration data, the actual flight data include redundant measurements of position and velocity from the global positioning system (GPS). The criteria derived from the blend of the GPS and INS accuracy was used to select valid trajectories for statistical analysis. The aerodynamic coefficients were derived from the selected trajectories by either direct extraction method based on the equations in dynamics, or by the inquiry of the pre-flight aerodynamic database. After the application of the proposed method to the case of the X-43A Hyper-X research vehicle, it was found that 1) there were consistent differences in the aerodynamic coefficients from the pre-flight aerodynamic database and post-flight analysis, 2) the pre-flight estimation of the pitching moment coefficients was significantly different from the post-flight analysis, 3) the type of distribution of the states from the Monte Carlo simulation were affected by that of the perturbation parameters, 4) the uncertainties in the pre-flight model were overestimated, 5) the range where the aerodynamic coefficients from the pre-flight aerodynamic database and post-flight analysis are in closest agreement is between Mach *.* and *.* and more data points may be needed between Mach * and ** in the pre-flight aerodynamic database, 6) selection criterion for valid trajectories from the Monte Carlo simulations was mostly driven by the horizontal velocity error, 7) the selection criterion must be based on reasonable model to ensure the validity of the statistics from the proposed method, and 8) the results from the proposed method applied to the two different flights with the identical geometry and similar flight profile were consistent.

In vivo measurement of aerodynamic weight support in freely flying birds

NASA Astrophysics Data System (ADS)

Lentink, David; Haselsteiner, Andreas; Ingersoll, Rivers

2014-11-01

Birds dynamically change the shape of their wing during the stroke to support their body weight aerodynamically. The wing is partially folded during the upstroke, which suggests that the upstroke of birds might not actively contribute to aerodynamic force production. This hypothesis is supported by the significant mass difference between the large pectoralis muscle that powers the down-stroke and the much smaller supracoracoideus that drives the upstroke. Previous works used indirect or incomplete techniques to measure the total force generated by bird wings ranging from muscle force, airflow, wing surface pressure, to detailed kinematics measurements coupled with bird mass-distribution models to derive net force through second derivatives. We have validated a new method that measures aerodynamic force in vivo time-resolved directly in freely flying birds which can resolve this question. The validation of the method, using independent force measurements on a quadcopter with pulsating thrust, show the aerodynamic force and impulse are measured within 2% accuracy and time-resolved. We demonstrate results for quad-copters and birds of similar weight and size. The method is scalable and can be applied to both engineered and natural flyers across taxa. The first author invented the method, the second and third authors validated the method and present results for quadcopters and birds.

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.

Aerodynamics model for a generic ASTOVL lift-fan aircraft

NASA Technical Reports Server (NTRS)

Birckelbaw, Lourdes G.; Mcneil, Walter E.; Wardwell, Douglas A.

1995-01-01

This report describes the aerodynamics model used in a simulation model of an advanced short takeoff and vertical landing (ASTOVL) lift-fan fighter aircraft. The simulation model was developed for use in piloted evaluations of transition and hover flight regimes, so that only low speed (M approximately 0.2) aerodynamics are included in the mathematical model. The aerodynamic model includes the power-off aerodynamic forces and moments and the propulsion system induced aerodynamic effects, including ground effects. The power-off aerodynamics data were generated using the U.S. Air Force Stability and Control Digital DATCOM program and a NASA Ames in-house graphics program called VORVIEW which allows the user to easily analyze arbitrary conceptual aircraft configurations using the VORLAX program. The jet-induced data were generated using the prediction methods of R. E. Kuhn et al., as referenced in this report.

Aerodynamic Flight-Test Results for the Adaptive Compliant Trailing Edge

NASA Technical Reports Server (NTRS)

Cumming, Stephen B.; Smith, Mark S.; Ali, Aliyah N.; Bui, Trong T.; Ellsworth, Joel C.; Garcia, Christian A.

2016-01-01

The aerodynamic effects of compliant flaps installed onto a modified Gulfstream III airplane were investigated. Analyses were performed prior to flight to predict the aerodynamic effects of the flap installation. Flight tests were conducted to gather both structural and aerodynamic data. The airplane was instrumented to collect vehicle aerodynamic data and wing pressure data. A leading-edge stagnation detection system was also installed. The data from these flights were analyzed and compared with predictions. The predictive tools compared well with flight data for small flap deflections, but differences between predictions and flight estimates were greater at larger deflections. This paper describes the methods used to examine the aerodynamics data from the flight tests and provides a discussion of the flight-test results in the areas of vehicle aerodynamics, wing sectional pressure coefficient profiles, and air data.

Fluid mechanics and solidification investigations in low-gravity environments

NASA Technical Reports Server (NTRS)

Fichtl, G. H.; Lundquist, C. A.; Naumann, R. J.

1980-01-01

Fluid mechanics of gases and liquids and solidification processes were investigated under microgravity conditions during Skylab and Apollo-Soyuz missions. Electromagnetic, acoustic, and aerodynamic levitation devices, drop tubes, aircraft parabolic flight trajectories, and vertical sounding rockets were developed for low-g simulation. The Spacelab 3 mission will be carried out in a gravity gradient flight attitude; analyses of sources of vehicle dynamic accelerations with associated g-levels and angular rates will produce results for future specific experiments.

Symposium on Numerical and Physical Aspects of Aerodynamic Flows

DTIC Science & Technology

1992-01-15

ReT/R. -1 wide range of Mach numbers including pressure gradient, a = - (a*) (10) surface roughness. surface heating and cooling . and surface 9 I + ReT...and specific heat at constant pressure, respectively. Favre equation is then obtained by contracting the Reynolds-stress decomposition is applied to...their near- wall behavior. In other words, if highly cooled -wall flows are to be predicted correctly, heat fluxes should be modeled e = 2ak + 4bky

Validation of a pair of computer codes for estimation and optimization of subsonic aerodynamic performance of simple hinged-flap systems for thin swept wings

NASA Technical Reports Server (NTRS)

Carlson, Harry W.; Darden, Christine M.

1988-01-01

Extensive correlations of computer code results with experimental data are employed to illustrate the use of linearized theory attached flow methods for the estimation and optimization of the aerodynamic performance of simple hinged flap systems. Use of attached flow methods is based on the premise that high levels of aerodynamic efficiency require a flow that is as nearly attached as circumstances permit. A variety of swept wing configurations are considered ranging from fighters to supersonic transports, all with leading- and trailing-edge flaps for enhancement of subsonic aerodynamic efficiency. The results indicate that linearized theory attached flow computer code methods provide a rational basis for the estimation and optimization of flap system aerodynamic performance at subsonic speeds. The analysis also indicates that vortex flap design is not an opposing approach but is closely related to attached flow design concepts. The successful vortex flap design actually suppresses the formation of detached vortices to produce a small vortex which is restricted almost entirely to the leading edge flap itself.

Validation of a computer code for analysis of subsonic aerodynamic performance of wings with flaps in combination with a canard or horizontal tail and an application to optimization

NASA Technical Reports Server (NTRS)

Carlson, Harry W.; Darden, Christine M.; Mann, Michael J.

1990-01-01

Extensive correlations of computer code results with experimental data are employed to illustrate the use of a linearized theory, attached flow method for the estimation and optimization of the longitudinal aerodynamic performance of wing-canard and wing-horizontal tail configurations which may employ simple hinged flap systems. Use of an attached flow method is based on the premise that high levels of aerodynamic efficiency require a flow that is as nearly attached as circumstances permit. The results indicate that linearized theory, attached flow, computer code methods (modified to include estimated attainable leading-edge thrust and an approximate representation of vortex forces) provide a rational basis for the estimation and optimization of aerodynamic performance at subsonic speeds below the drag rise Mach number. Generally, good prediction of aerodynamic performance, as measured by the suction parameter, can be expected for near optimum combinations of canard or horizontal tail incidence and leading- and trailing-edge flap deflections at a given lift coefficient (conditions which tend to produce a predominantly attached flow).

Laryngeal Aerodynamics in Healthy Older Adults and Adults with Parkinson's Disease

ERIC Educational Resources Information Center

Matheron, Deborah; Stathopoulos, Elaine T.; Huber, Jessica E.; Sussman, Joan E.

2017-01-01

Purpose: The present study compared laryngeal aerodynamic function of healthy older adults (HOA) to adults with Parkinson's disease (PD) while speaking at a comfortable and increased vocal intensity. Method: Laryngeal aerodynamic measures (subglottal pressure, peak-to-peak flow, minimum flow, and open quotient [OQ]) were compared between HOAs and…

Comparison of Theodorsen's Unsteady Aerodynamic Forces with Doublet Lattice Generalized Aerodynamic Forces

NASA Technical Reports Server (NTRS)

Perry, Boyd, III

2017-01-01

This paper identifies the unsteady aerodynamic forces and moments for a typical section contained in the NACA Report No. 496, "General Theory of Aerodynamic Instability and the Mechanism of Flutter," by Theodore Theodorsen. These quantities are named Theodorsen's aerodynamic forces (TAFs). The TAFs are compared to the generalized aerodynamic forces (GAFs) for a very high aspect ratio wing (AR = 20) at zero Mach number computed by the doublet lattice method. Agreement between TAFs and GAFs is very-good-to-excellent. The paper also reveals that simple proportionality relationships that are known to exist between the real parts of some GAFs and the imaginary parts of others also hold for the real and imaginary parts of the corresponding TAFs.

Development of a nonlinear vortex method. [steady and unsteady aerodynamic loads of highly sweptback wings

NASA Technical Reports Server (NTRS)

Kandil, O. A.

1981-01-01

Progress is reported in the development of reliable nonlinear vortex methods for predicting the steady and unsteady aerodynamic loads of highly sweptback wings at large angles of attack. Abstracts of the papers, talks, and theses produced through this research are included. The modified nonlinear discrete vortex method and the nonlinear hybrid vortex method are highlighted.

Assessment of the Derivative-Moment Transformation method for unsteady-load estimation

NASA Astrophysics Data System (ADS)

Mohebbian, Ali; Rival, David

2011-11-01

It is often difficult, if not impossible, to measure the aerodynamic or hydrodynamic forces on a moving body. For this reason, a classical control-volume technique is typically applied to extract the unsteady forces instead. However, measuring the acceleration term within the volume of interest using PIV can be limited by optical access, reflections as well as shadows. Therefore in this study an alternative approach, termed the Derivative-Moment Transformation (DMT) method, is introduced and tested on a synthetic data set produced using numerical simulations. The test case involves the unsteady loading of a flat plate in a two-dimensional, laminar periodic gust. The results suggest that the DMT method can accurately predict the acceleration term so long as appropriate spatial and temporal resolutions are maintained. The major deficiency was found to be the determination of pressure in the wake. The effect of control-volume size was investigated suggesting that smaller domains work best by minimizing the associated error with the pressure field. When increasing the control-volume size, the number of calculations necessary for the pressure-gradient integration increases, in turn substantially increasing the error propagation.

A High-Lift Building Block Flow: Turbulent Boundary Layer Relaminarization

NASA Technical Reports Server (NTRS)

Bourassa, Corey; Thomas, Flint O.; Nelson, Robert C.

2001-01-01

A working wind tunnel test facility has been constructed at the University of Notre Dame's Hessert Center. The relaminarization test facility has been constructed in the 1.5m x 1.5m (5ft x 5 ft) atmospheric wind tunnel and generates a Re(theta)=4694 turbulent boundary layer in nominally zero-pressure gradient before it is exposed to the Case #1 pressure gradient (K approximately equal to 4.2 x 10(exp -6), which is believed to be sufficient to achieve relaminarization. Future work to be conducted will include measuring the response of the turbulent boundary layer to the favorable pressure gradients created in the test facility and documenting this response in order to understand the underlying flow physics responsible for relaminarization. It is the goal of this research to have a better understanding of accelerated turbulent boundary layers which will aid in the development of future flow diagnostic utilities to be implemented in applied aerodynamic research.

Measurements of Aerodynamic Damping in the MIT Transonic Rotor

NASA Technical Reports Server (NTRS)

Crawley, E. F.

1981-01-01

A method was developed and demonstrated for the direct measurement of aerodynamic forcing and aerodynamic damping of a transonic compressor. The method is based on the inverse solution of the structural dynamic equations of motion of the blade disk system in order to determine the forces acting on the system. The disturbing and damping forces acting on a given blade are determined if the equations of motion are expressed in individual blade coordinates. If the structural dynamic equations are transformed to multiblade coordinates, the damping can be measured for blade disk modes, and related to a reduced frequency and interblade phase angle. In order to measure the aerodynamic damping in this way, the free response to a known excitation is studied.

Physically weighted approximations of unsteady aerodynamic forces using the minimum-state method

NASA Technical Reports Server (NTRS)

Karpel, Mordechay; Hoadley, Sherwood Tiffany

1991-01-01

The Minimum-State Method for rational approximation of unsteady aerodynamic force coefficient matrices, modified to allow physical weighting of the tabulated aerodynamic data, is presented. The approximation formula and the associated time-domain, state-space, open-loop equations of motion are given, and the numerical procedure for calculating the approximation matrices, with weighted data and with various equality constraints are described. Two data weighting options are presented. The first weighting is for normalizing the aerodynamic data to maximum unit value of each aerodynamic coefficient. The second weighting is one in which each tabulated coefficient, at each reduced frequency value, is weighted according to the effect of an incremental error of this coefficient on aeroelastic characteristics of the system. This weighting yields a better fit of the more important terms, at the expense of less important ones. The resulting approximate yields a relatively low number of aerodynamic lag states in the subsequent state-space model. The formulation forms the basis of the MIST computer program which is written in FORTRAN for use on the MicroVAX computer and interfaces with NASA's Interaction of Structures, Aerodynamics and Controls (ISAC) computer program. The program structure, capabilities and interfaces are outlined in the appendices, and a numerical example which utilizes Rockwell's Active Flexible Wing (AFW) model is given and discussed.

Special methods for aerodynamic-moment calculations from parachute FSI modeling

NASA Astrophysics Data System (ADS)

Takizawa, Kenji; Tezduyar, Tayfun E.; Boswell, Cody; Tsutsui, Yuki; Montel, Kenneth

2015-06-01

The space-time fluid-structure interaction (STFSI) methods for 3D parachute modeling are now at a level where they can bring reliable, practical analysis to some of the most complex parachute systems, such as spacecraft parachutes. The methods include the Deforming-Spatial-Domain/Stabilized ST method as the core computational technology, and a good number of special FSI methods targeting parachutes. Evaluating the stability characteristics of a parachute based on how the aerodynamic moment varies as a function of the angle of attack is one of the practical analyses that reliable parachute FSI modeling can deliver. We describe the special FSI methods we developed for this specific purpose and present the aerodynamic-moment data obtained from FSI modeling of NASA Orion spacecraft parachutes and Japan Aerospace Exploration Agency (JAXA) subscale parachutes.

Simultaneous Aerodynamic and Structural Design Optimization (SASDO) for a 3-D Wing

NASA Technical Reports Server (NTRS)

Gumbert, Clyde R.; Hou, Gene J.-W.; Newman, Perry A.

2001-01-01

The formulation and implementation of an optimization method called Simultaneous Aerodynamic and Structural Design Optimization (SASDO) is shown as an extension of the Simultaneous Aerodynamic Analysis and Design Optimization (SAADO) method. It is extended by the inclusion of structure element sizing parameters as design variables and Finite Element Method (FEM) analysis responses as constraints. The method aims to reduce the computational expense. incurred in performing shape and sizing optimization using state-of-the-art Computational Fluid Dynamics (CFD) flow analysis, FEM structural analysis and sensitivity analysis tools. SASDO is applied to a simple. isolated, 3-D wing in inviscid flow. Results show that the method finds the saine local optimum as a conventional optimization method with some reduction in the computational cost and without significant modifications; to the analysis tools.

Study of wavefront error and polarization of a side mounted infrared window

NASA Astrophysics Data System (ADS)

Liu, Jiaguo; Li, Lin; Hu, Xinqi; Yu, Xin

2008-03-01

The wavefront error and polarization of a side mounted infrared window made of ZnS are studied. The Infrared windows suffer from temperature gradient and stress during their launch process. Generally, the gradient in temperature changes the refractive index of the material whereas stress produces deformation and birefringence. In this paper, a thermal finite element analysis (FEA) of an IR window is presented. For this purpose, we employed an FEA program Ansys to obtain the time-varying temperature field. The deformation and stress of the window are derived from a structural FEA with the aerodynamic force and the temperature field previously obtained as being the loads. The deformation, temperature field, stress field, ray tracing and Jones Calculus are used to calculate the wavefront error and the change of polarization state.

Aerodynamic Modeling of Oscillating Wing in Hypersonic Flow: a Numerical Study

NASA Astrophysics Data System (ADS)

Zhu, Jian; Hou, Ying-Yu; Ji, Chen; Liu, Zi-Qiang

2016-06-01

Various approximations to unsteady aerodynamics are examined for the unsteady aerodynamic force of a pitching thin double wedge airfoil in hypersonic flow. Results of piston theory, Van Dyke’s second-order theory, Newtonian impact theory, and CFD method are compared in the same motion and Mach number effects. The results indicate that, for this thin double wedge airfoil, Newtonian impact theory is not suitable for these Mach number, while piston theory and Van Dyke’s second-order theory are in good agreement with CFD method for Ma<7.

Experimental determination of damping of plate vibrations in a viscous fluid

NASA Astrophysics Data System (ADS)

Egorov, A. G.; Kamalutdinov, A. M.; Nuriev, A. N.; Paimushin, V. N.

2017-05-01

A method of determining the aerodynamic-drag coefficient of flat vibrating plates from the vibrogram of free damping vibrations of cantilever-fixed duralumin samples has been developed. From the results of our experiments, simple approximating formulas determining the decrement of damping vibrations and the aerodynamic-drag coefficient through the dimensionless vibration amplitude and the Stokes parameter are proposed. The approach developed in this study for determining the aerodynamic-drag coefficient of a vibrating plate can be a useful alternative to purely hydrodynamic methods of finding the drag of vibrating solids.

Semiempirical method for prediction of aerodynamic forces and moments on a steadily spinning light airplane

NASA Technical Reports Server (NTRS)

Pamadi, Bandu N.; Taylor, Lawrence W., Jr.

1987-01-01

A semi-empirical method is presented for the estimation of aerodynamic forces and moments acting on a steadily spinning (rotating) light airplane. The airplane is divided into wing, body, and tail surfaces. The effect of power is ignored. The strip theory is employed for each component of the spinning airplane to determine its contribution to the total aerodynamic coefficients. Then, increments to some of the coefficients which account for centrifugal effect are estimated. The results are compared to spin tunnel rotary balance test data.

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.

Airfoil structure

DOEpatents

Frey, Gary A.; Twardochleb, Christopher Z.

1998-01-01

Past airfoil configurations have been used to improve aerodynamic performance and engine efficiencies. The present airfoil configuration further increases component life and reduces maintenance by reducing internal stress within the airfoil itself. The airfoil includes a chord and a span. Each of the chord and the span has a bow being summed to form a generally "C" configuration of the airfoil. The generally "C" configuration includes a compound bow in which internal stresses resulting from a thermal temperature gradient are reduced. The structural configuration reduces internal stresses resulting from thermal expansion.

Airfoil structure

DOEpatents

Frey, G.A.; Twardochleb, C.Z.

1998-01-13

Past airfoil configurations have been used to improve aerodynamic performance and engine efficiencies. The present airfoil configuration further increases component life and reduces maintenance by reducing internal stress within the airfoil itself. The airfoil includes a chord and a span. Each of the chord and the span has a bow being summed to form a generally ``C`` configuration of the airfoil. The generally ``C`` configuration includes a compound bow in which internal stresses resulting from a thermal temperature gradient are reduced. The structural configuration reduces internal stresses resulting from thermal expansion. 6 figs.

Proceedings of the Conference on Improvement of Aerodynamic Performance through Boundary Layer Control and High Lift Systems Held at the Fluid Dynamics Panel Symposium in Brussels, Belgium on 21-23 May 1984.

DTIC Science & Technology

1984-08-01

transversal eat dirigi veta ia partie concave de ia ligne de courant ext~rieurs (figure 2). L~ e gradient de pression longitudinal change ensuite de signe, mais...U-° II ’a. .- -- £ V -.. .. ,.% .: , -. _ . . .. -. . . .... .. . . . . . .. .. . .... ... ,.. ..44*- , . . .* . .. -...... . . . . . -- ,, e . . e ...8217 ," . • " ." ,, ,, o ’_; e , ,., . , , " • , . . . . . . ,*. . . . . . . .*’ # - * -~.- ~-**-JJI---w ~ ’ AGARD-CP-365 I N-. NORTH ATLANTIC

Deployment and Intelligent Nanosatellite Operations Colorado Final Technical Report

DTIC Science & Technology

2006-09-28

environmental factors will cause disturbance torques during orbit around the Earth . These factors are solar radiation pressure from the sun , aerodynamic...software. The 3- axis sensing of the magnetometer allows a vector the B- field of the Earth to be sensed. Geopack 2003 then can be utilized with the orbit ...gradient torque can be represented as the following: g, ; 3wo21 Eq. 2-11 where ow is the angular velocity of the spacecraft as it orbits the earth . DINO’s

Gravitational orientation of the orbital complex, Salyut-6--Soyuz

NASA Technical Reports Server (NTRS)

Grecho, G. M.; Sarychev, V. A.; Legostayev, V. P.; Sazonov, V. V.; Gansvind, I. N.

1983-01-01

A simple mathematical model is proposed for the Salyut-6-Soyuz orbital complex motion with respect to the center of mass under the one-axis gravity-gradient orientation regime. This model was used for processing the measurements of the orbital complex motion parameters when the above orientation region was implemented. Some actual satellite motions are simulated and the satellite's aerodynamic parameters are determined. Estimates are obtained for the accuracy of measurements as well as that of the mathematical model.

Domain decomposition methods in aerodynamics

NASA Technical Reports Server (NTRS)

Venkatakrishnan, V.; Saltz, Joel

1990-01-01

Compressible Euler equations are solved for two-dimensional problems by a preconditioned conjugate gradient-like technique. An approximate Riemann solver is used to compute the numerical fluxes to second order accuracy in space. Two ways to achieve parallelism are tested, one which makes use of parallelism inherent in triangular solves and the other which employs domain decomposition techniques. The vectorization/parallelism in triangular solves is realized by the use of a recording technique called wavefront ordering. This process involves the interpretation of the triangular matrix as a directed graph and the analysis of the data dependencies. It is noted that the factorization can also be done in parallel with the wave front ordering. The performances of two ways of partitioning the domain, strips and slabs, are compared. Results on Cray YMP are reported for an inviscid transonic test case. The performances of linear algebra kernels are also reported.

Basis Function Approximation of Transonic Aerodynamic Influence Coefficient Matrix

NASA Technical Reports Server (NTRS)

Li, Wesley Waisang; Pak, Chan-Gi

2010-01-01

A technique for approximating the modal aerodynamic influence coefficients [AIC] matrices by using basis functions has been developed and validated. An application of the resulting approximated modal AIC matrix for a flutter analysis in transonic speed regime has been demonstrated. This methodology can be applied to the unsteady subsonic, transonic and supersonic aerodynamics. The method requires the unsteady aerodynamics in frequency-domain. The flutter solution can be found by the classic methods, such as rational function approximation, k, p-k, p, root-locus et cetera. The unsteady aeroelastic analysis for design optimization using unsteady transonic aerodynamic approximation is being demonstrated using the ZAERO(TradeMark) flutter solver (ZONA Technology Incorporated, Scottsdale, Arizona). The technique presented has been shown to offer consistent flutter speed prediction on an aerostructures test wing [ATW] 2 configuration with negligible loss in precision in transonic speed regime. These results may have practical significance in the analysis of aircraft aeroelastic calculation and could lead to a more efficient design optimization cycle

Aerothermoelastic response analysis for C/SiC panel of ceramic matrix composite shingle thermal protection system

NASA Astrophysics Data System (ADS)

Huo, Lin; Cheng, Xing-Hua; Yang, Tao

2015-05-01

This paper presents a study of aerothermoelastic response of a C/SiC panel, which is a primary structure for ceramic matrix composite shingle thermal protection system for hypersonic vehicles. It is based on a three dimensional thermal protection shingle panel on a quasi-waverider vehicle model. Firstly, the Thin Shock Layer and piston theory are adopted to compute the aerodynamic pressure of rigid body and deformable body, and a series of engineering methods are used to compute the aerodynamic heating. Then an aerothermoelastic loosely-coupled time marching strategy and self-adapting aerodynamic heating time step are developed to analyze the aerothermoelastic response of the panel, with an aerodynamic heating and temperature field coupling parameter selection method being adopted to increase the efficiency. Finally, a few revealing conclusions are reached by analyzing how coupling at different degrees influences the quasi-static aerothermoelastic response of the panel and how aerodynamic pressure of rigid body time step influences the quasi-static aerothermoelastic response on a glide trajectory.

Methods for extracting aerodynamic accelerations from Orbiter High Resolution Accelerometer Package flight data

NASA Technical Reports Server (NTRS)

Thompson, J. M.; Russell, J. W.; Blanchard, R. C.

1987-01-01

This report presents a process for extracting the aerodynamic accelerations of the Shuttle Orbiter Vehicle from the High Resolution Accelerometer Package (HiRAP) flight data during reentry. The methods for obtaining low-level aerodynamic accelerations, principally in the rarefied flow regime, are applied to 10 Orbiter flights. The extraction process is presented using data obtained from Space Transportation System Flight 32 (Mission 61-C) as a typical example. This process involves correcting the HiRAP measurements for the effects of temperature bias and instrument offset from the Orbiter center of gravity, and removing acceleration data during times they are affected by thruster firings. The corrected data are then made continuous and smooth and are further enhanced by refining the temperature bias correction and removing effects of the auxiliary power unit actuation. The resulting data are the current best estimate of the Orbiter aerodynamic accelerations during reentry and will be used for further analyses of the Orbiter aerodynamics and the upper atmosphere characteristics.

Individual Monitoring of Vocal Effort With Relative Fundamental Frequency: Relationships With Aerodynamics and Listener Perception

PubMed Central

Michener, Carolyn M.; Eadie, Tanya L.; Stepp, Cara E.

2015-01-01

Purpose The acoustic measure relative fundamental frequency (RFF) was investigated as a potential objective measure to track variations in vocal effort within and across individuals. Method Twelve speakers with healthy voices created purposeful modulations in their vocal effort during speech tasks. RFF and an aerodynamic measure of vocal effort, the ratio of sound pressure level to subglottal pressure level, were estimated from the aerodynamic and acoustic signals. Twelve listeners also judged the speech samples for vocal effort using the visual sort and rate method. Results Relationships between RFF and both the aerodynamic and perceptual measures of vocal effort were weak across speakers (R2 = .06–.26). Within speakers, relationships were variable but much stronger on average (R2 = .45–.56). Conclusions RFF showed stronger relationships between both the aerodynamic and perceptual measures of vocal effort when examined within individuals versus across individuals. Future work is necessary to establish these relationships in individuals with voice disorders across the therapeutic process. PMID:25675090

Numerical method to compute acoustic scattering effect of a moving source.

PubMed

Song, Hao; Yi, Mingxu; Huang, Jun; Pan, Yalin; Liu, Dawei

2016-01-01

In this paper, the aerodynamic characteristic of a ducted tail rotor in hover has been numerically studied using CFD method. An analytical time domain formulation based on Ffowcs Williams-Hawkings (FW-H) equation is derived for the prediction of the acoustic velocity field and used as Neumann boundary condition on a rigid scattering surface. In order to predict the aerodynamic noise, a hybrid method combing computational aeroacoustics with an acoustic thin-body boundary element method has been proposed. The aerodynamic results and the calculated sound pressure levels (SPLs) are compared with the known method for validation. Simulation results show that the duct can change the value of SPLs and the sound directivity. Compared with the isolate tail rotor, the SPLs of the ducted tail rotor are smaller at certain azimuth.

Current status of computational methods for transonic unsteady aerodynamics and aeroelastic applications

NASA Technical Reports Server (NTRS)

Edwards, John W.; Malone, John B.

1992-01-01

The current status of computational methods for unsteady aerodynamics and aeroelasticity is reviewed. The key features of challenging aeroelastic applications are discussed in terms of the flowfield state: low-angle high speed flows and high-angle vortex-dominated flows. The critical role played by viscous effects in determining aeroelastic stability for conditions of incipient flow separation is stressed. The need for a variety of flow modeling tools, from linear formulations to implementations of the Navier-Stokes equations, is emphasized. Estimates of computer run times for flutter calculations using several computational methods are given. Applications of these methods for unsteady aerodynamic and transonic flutter calculations for airfoils, wings, and configurations are summarized. Finally, recommendations are made concerning future research directions.

Aerodynamic characterisation and trajectory simulations for the Ariane-5 booster recovery system

NASA Astrophysics Data System (ADS)

Meiboom, F. P.

One of the most critical aspects of the early phases of the development of the Ariane-5 booster recovery system was the determination of the behavior of the booster during its atmospheric reentry, since this behavior determines the start conditions for the parachute system elements. A combination of wind-tunnel tests (subsonic and supersonic) and analytical methods was applied to define the aerodynamic characteristics of the booster. This aerodynamic characterization in combination with information of the ascent trajectory, atmospheric properties and booster mass and inertia were used as input for the 6-DOF trajectory simulations of the vehicle. Uncertainties in aerodynamic properties and deviations in atmospheric and booster properties were incorporated to define the range of initial conditions for the parachute system, utilizing stochastic (Monte-Carlo) methods.

Successive smoothing algorithm for constructing the semiempirical model developed at ONERA to predict unsteady aerodynamic forces. [aeroelasticity in helicopters

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.

Missile Aerodynamics

DTIC Science & Technology

1979-02-01

aimed to emphasize these differences in the aerodynamic design features of both guided and unguided weapons. In addition to treating the component parts...the subject. Lectures generally started with a review of fundamentals and paid particular attention to practical methods of estimation and design and...George G Brebner Aerodynamics Department Royal Aircraft Establishment Farnborough, Hants, GU14 6TD, UK SUMMARY The differences in design objectives and

Sonic boom generated by a slender body aerodynamically shaded by a disk spike

NASA Astrophysics Data System (ADS)

Potapkin, A. V.; Moskvichev, D. Yu.

2018-03-01

The sonic boom generated by a slender body of revolution aerodynamically shaded by another body is numerically investigated. The aerodynamic shadow is created by a disk placed upstream of the slender body across a supersonic free-stream flow. The disk size and its position upstream of the body are chosen in such a way that the aerodynamically shaded flow is quasi-stationary. A combined method of phantom bodies is used for sonic boom calculations. The method is tested by calculating the sonic boom generated by a blunted body and comparing the results with experimental investigations of the sonic boom generated by spheres of various diameters in ballistic ranges and wind tunnels. The test calculations show that the method of phantom bodies is applicable for calculating far-field parameters of shock waves generated by both slender and blunted bodies. A possibility of reducing the shock wave intensity in the far field by means of the formation of the aerodynamic shadow behind the disk placed upstream of the body is estimated. The calculations are performed for the incoming flow with the Mach number equal to 2. The effect of the disk size on the sonic boom level is calculated.

External-Compression Supersonic Inlet Design Code

NASA Technical Reports Server (NTRS)

Slater, John W.

2011-01-01

A computer code named SUPIN has been developed to perform aerodynamic design and analysis of external-compression, supersonic inlets. The baseline set of inlets include axisymmetric pitot, two-dimensional single-duct, axisymmetric outward-turning, and two-dimensional bifurcated-duct inlets. The aerodynamic methods are based on low-fidelity analytical and numerical procedures. The geometric methods are based on planar geometry elements. SUPIN has three modes of operation: 1) generate the inlet geometry from a explicit set of geometry information, 2) size and design the inlet geometry and analyze the aerodynamic performance, and 3) compute the aerodynamic performance of a specified inlet geometry. The aerodynamic performance quantities includes inlet flow rates, total pressure recovery, and drag. The geometry output from SUPIN includes inlet dimensions, cross-sectional areas, coordinates of planar profiles, and surface grids suitable for input to grid generators for analysis by computational fluid dynamics (CFD) methods. The input data file for SUPIN and the output file from SUPIN are text (ASCII) files. The surface grid files are output as formatted Plot3D or stereolithography (STL) files. SUPIN executes in batch mode and is available as a Microsoft Windows executable and Fortran95 source code with a makefile for Linux.

Influence of unsteady aerodynamics on driving dynamics of passenger cars

NASA Astrophysics Data System (ADS)

Huemer, Jakob; Stickel, Thomas; Sagan, Erich; Schwarz, Martin; Wall, Wolfgang A.

2014-11-01

Recent approaches towards numerical investigations with computational fluid dynamics methods on unsteady aerodynamic loads of passenger cars identified major differences compared with steady-state aerodynamic excitations. Furthermore, innovative vehicle concepts such as electric-vehicles or hybrid drives further challenge the basic layout of passenger cars. Therefore, the relevance of unsteady aerodynamic loads on cross-wind stability of changing basic vehicle architectures should be analysed. In order to assure and improve handling and ride characteristics at high velocity of the actual range of vehicle layouts, the influence of unsteady excitations on the vehicle response was investigated. For this purpose, a simulation of the vehicle dynamics through multi-body simulation was used. The impact of certain unsteady aerodynamic load characteristics on the vehicle response was quantified and key factors were identified. Through a series of driving simulator tests, the identified differences in the vehicle response were evaluated regarding their significance on the subjective driver perception of cross-wind stability. Relevant criteria for the subjective driver assessment of the vehicle response were identified. As a consequence, a design method for the basic layout of passenger cars and chassis towards unsteady aerodynamic excitations was defined.

High-speed aerodynamic design of space vehicle and required hypersonic wind tunnel facilities

NASA Astrophysics Data System (ADS)

Sakakibara, Seizou; Hozumi, Kouichi; Soga, Kunio; Nomura, Shigeaki

Problems associated with the aerodynamic design of space vehicles with emphasis of the role of hypersonic wind tunnel facilities in the development of the vehicle are considered. At first, to identify wind tunnel and computational fluid dynamics (CFD) requirements, operational environments are postulated for hypervelocity vehicles. Typical flight corridors are shown with the associated flow density: real gas effects, low density flow, and non-equilibrium flow. Based on an evaluation of these flight regimes and consideration of the operational requirements, the wind tunnel testing requirements for the aerodynamic design are examined. Then, the aerodynamic design logic and optimization techniques to develop and refine the configurations in a traditional phased approach based on the programmatic design of space vehicle are considered. Current design methodology for the determination of aerodynamic characteristics for designing the space vehicle, i.e., (1) ground test data, (2) numerical flow field solutions and (3) flight test data, are also discussed. Based on these considerations and by identifying capabilities and limits of experimental and computational methods, the role of a large conventional hypersonic wind tunnel and the high enthalpy tunnel and the interrelationship of the wind tunnels and CFD methods in actual aerodynamic design and analysis are discussed.

Aerodynamics Via Acoustics: Application of Acoustic Formulas for Aerodynamic Calculations

NASA Technical Reports Server (NTRS)

Farassat, F.; Myers, M. K.

1986-01-01

Prediction of aerodynamic loads on bodies in arbitrary motion is considered from an acoustic point of view, i.e., in a frame of reference fixed in the undisturbed medium. An inhomogeneous wave equation which governs the disturbance pressure is constructed and solved formally using generalized function theory. When the observer is located on the moving body surface there results a singular linear integral equation for surface pressure. Two different methods for obtaining such equations are discussed. Both steady and unsteady aerodynamic calculations are considered. Two examples are presented, the more important being an application to propeller aerodynamics. Of particular interest for numerical applications is the analytical behavior of the kernel functions in the various integral equations.

Flow pattern similarities in the near wake of three bird species suggest a common role for unsteady aerodynamic effects in lift generation

PubMed Central

Krishnan, Krishnamoorthy; Ben-Gida, Hadar; Kirchhefer, Adam J.; Kopp, Gregory A.; Guglielmo, Christopher G.

2017-01-01

Analysis of the aerodynamics of flapping wings has yielded a general understanding of how birds generate lift and thrust during flight. However, the role of unsteady aerodynamics in avian flight due to the flapping motion still holds open questions in respect to performance and efficiency. We studied the flight of three distinctive bird species: western sandpiper (Calidris mauri), European starling (Sturnus vulgaris) and American robin (Turdus migratorius) using long-duration, time-resolved particle image velocimetry, to better characterize and advance our understanding of how birds use unsteady flow features to enhance their aerodynamic performances during flapping flight. We show that during transitions between downstroke and upstroke phases of the wing cycle, the near wake-flow structures vary and generate unique sets of vortices. These structures appear as quadruple layers of concentrated vorticity aligned at an angle with respect to the horizon (named ‘double branch’). They occur where the circulation gradient changes sign, which implies that the forces exerted by the flapping wings of birds are modified during the transition phases. The flow patterns are similar in (non-dimensional) size and magnitude for the different birds suggesting that there are common mechanisms operating during flapping flight across species. These flow patterns occur at the same phase where drag reduction of about 5% per cycle and lift enhancement were observed in our prior studies. We propose that these flow structures should be considered in wake flow models that seek to account for the contribution of unsteady flow to lift and drag. PMID:28163881

Free Wake Techniques for Rotor Aerodynamic Analysis. Volume 1: Summary of Results and Background Theory

NASA Technical Reports Server (NTRS)

Miller, R. H.

1982-01-01

Results obtained during the development of a consistent aerodynamic theory for rotors in hovering flight are discussed. Methods of aerodynamic analysis were developed which are adequate for general design purposes until such time as more elaborate solutions become available, in particular solutions which include real fluids effects. Several problems were encountered in the course of this development, and many remain to be solved, however it is felt that a better understanding of the aerodynamic phenomena involved was obtained. Remaining uncertainties are discussed.

Unsteady Aerodynamic Force Sensing from Strain Data

NASA Technical Reports Server (NTRS)

Pak, Chan-Gi

2017-01-01

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.

Prediction of Hyper-X Stage Separation Aerodynamics Using CFD

NASA Technical Reports Server (NTRS)

Buning, Pieter G.; Wong, Tin-Chee; Dilley, Arthur D.; Pao, Jenn L.

2000-01-01

The NASA X-43 "Hyper-X" hypersonic research vehicle will be boosted to a Mach 7 flight test condition mounted on the nose of an Orbital Sciences Pegasus launch vehicle. The separation of the research vehicle from the Pegasus presents some unique aerodynamic problems, for which computational fluid dynamics has played a role in the analysis. This paper describes the use of several CFD methods for investigating the aerodynamics of the research and launch vehicles in close proximity. Specifically addressed are unsteady effects, aerodynamic database extrapolation, and differences between wind tunnel and flight environments.

The space shuttle ascent vehicle aerodynamic challenges configuration design and data base development

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.

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.

1999 NASA High-Speed Research Program Aerodynamic Performance Workshop. Volume 2; High Lift

NASA Technical Reports Server (NTRS)

Hahne, David E. (Editor)

1999-01-01

NASA's High-Speed Research Program sponsored the 1999 Aerodynamic Performance Technical Review on February 8-12, 1999 in Anaheim, California. The review was designed to bring together NASA and industry High-Speed Civil Transport (HSCT) Aerodynamic Performance technology development participants in the areas of Configuration Aerodynamics (transonic and supersonic cruise drag prediction and minimization), High Lift, and Flight Controls. The review objectives were to (1) report the progress and status of HSCT aerodynamic performance technology development; (2) disseminate this technology within the appropriate technical communities; and (3) promote synergy among die scientists and engineers working on HSCT aerodynamics. In particular, single and midpoint optimized HSCT configurations, HSCT high-lift system performance predictions, and HSCT simulation results were presented, along with executive summaries for all the Aerodynamic Performance technology areas. The HSR Aerodynamic Performance Technical Review was held simultaneously with the annual review of the following airframe technology areas: Materials and Structures, Environmental Impact, Flight Deck, and Technology Integration. Thus, a fourth objective of the Review was to promote synergy between the Aerodynamic Performance technology area and the other technology areas of the HSR Program. This Volume 2/Part 2 publication covers the tools and methods development session.

Wing-section optimization for supersonic viscous flow

NASA Technical Reports Server (NTRS)

Item, Cem C.; Baysal, Oktay (Editor)

1995-01-01

To improve the shape of a supersonic wing, an automated method that also includes higher fidelity to the flow physics is desirable. With this impetus, an aerodynamic optimization methodology incorporating thin-layer Navier-Stokes equations and sensitivity analysis had been previously developed. Prior to embarking upon the wind design task, the present investigation concentrated on testing the feasibility of the methodology, and the identification of adequate problem formulations, by defining two-dimensional, cost-effective test cases. Starting with two distinctly different initial airfoils, two independent shape optimizations resulted in shapes with similar features: slightly cambered, parabolic profiles with sharp leading- and trailing-edges. Secondly, the normal section to the subsonic portion of the leading edge, which had a high normal angle-of-attack, was considered. The optimization resulted in a shape with twist and camber which eliminated the adverse pressure gradient, hence, exploiting the leading-edge thrust. The wing section shapes obtained in all the test cases had the features predicted by previous studies. Therefore, it was concluded that the flowfield analyses and sensitivity coefficients were computed and fed to the present gradient-based optimizer correctly. Also, as a result of the present two-dimensional study, suggestions were made for the problem formulations which should contribute to an effective wing shape optimization.

Comparison of Computational Approaches for Rapid Aerodynamic Assessment of Small UAVs

NASA Technical Reports Server (NTRS)

Shafer, Theresa C.; Lynch, C. Eric; Viken, Sally A.; Favaregh, Noah; Zeune, Cale; Williams, Nathan; Dansie, Jonathan

2014-01-01

Computational Fluid Dynamic (CFD) methods were used to determine the basic aerodynamic, performance, and stability and control characteristics of the unmanned air vehicle (UAV), Kahu. Accurate and timely prediction of the aerodynamic characteristics of small UAVs is an essential part of military system acquisition and air-worthiness evaluations. The forces and moments of the UAV were predicted using a variety of analytical methods for a range of configurations and conditions. The methods included Navier Stokes (N-S) flow solvers (USM3D, Kestrel and Cobalt) that take days to set up and hours to converge on a single solution; potential flow methods (PMARC, LSAERO, and XFLR5) that take hours to set up and minutes to compute; empirical methods (Datcom) that involve table lookups and produce a solution quickly; and handbook calculations. A preliminary aerodynamic database can be developed very efficiently by using a combination of computational tools. The database can be generated with low-order and empirical methods in linear regions, then replacing or adjusting the data as predictions from higher order methods are obtained. A comparison of results from all the data sources as well as experimental data obtained from a wind-tunnel test will be shown and the methods will be evaluated on their utility during each portion of the flight envelope.

Evaluation of aerodynamic derivatives from a magnetic balance system

NASA Technical Reports Server (NTRS)

Raghunath, B. S.; Parker, H. M.

1972-01-01

The dynamic testing of a model in the University of Virginia cold magnetic balance wind-tunnel facility is expected to consist of measurements of the balance forces and moments, and the observation of the essentially six degree of freedom motion of the model. The aerodynamic derivatives of the model are to be evaluated from these observations. The basic feasibility of extracting aerodynamic information from the observation of a model which is executing transient, complex, multi-degree of freedom motion is demonstrated. It is considered significant that, though the problem treated here involves only linear aerodynamics, the methods used are capable of handling a very large class of aerodynamic nonlinearities. The basic considerations include the effect of noise in the data on the accuracy of the extracted information. Relationships between noise level and the accuracy of the evaluated aerodynamic derivatives are presented.

Prediction of Aerodynamic Coefficient using Genetic Algorithm Optimized Neural Network for Sparse Data

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 coefficients to an accuracy of 110% . In our problem, we would like to get an optimized neural network architecture and minimum data set. This has been accomplished within 500 training cycles of a neural network. After removing training pairs (outliers), the GA has produced much better results. The neural network constructed is a feed forward neural network with a back propagation learning mechanism. The main goal has been to free the network design process from constraints of human biases, and to discover better forms of neural network architectures. The automation of the network architecture search by genetic algorithms seems to have been the best way to achieve this goal.

Aircraft Flight Envelope Determination using Upset Detection and Physical Modeling Methods

NASA Technical Reports Server (NTRS)

Keller, Jeffrey D.; McKillip, Robert M. Jr.; Kim, Singwan

2009-01-01

The development of flight control systems to enhance aircraft safety during periods of vehicle impairment or degraded operations has been the focus of extensive work in recent years. Conditions adversely affecting aircraft flight operations and safety may result from a number of causes, including environmental disturbances, degraded flight operations, and aerodynamic upsets. To enhance the effectiveness of adaptive and envelope limiting controls systems, it is desirable to examine methods for identifying the occurrence of anomalous conditions and for assessing the impact of these conditions on the aircraft operational limits. This paper describes initial work performed toward this end, examining the use of fault detection methods applied to the aircraft for aerodynamic performance degradation identification and model-based methods for envelope prediction. Results are presented in which a model-based fault detection filter is applied to the identification of aircraft control surface and stall departure failures/upsets. This application is supported by a distributed loading aerodynamics formulation for the flight dynamics system reference model. Extensions for estimating the flight envelope due to generalized aerodynamic performance degradation are also described.

Aerodynamic Validation of Emerging Projectile Configurations

DTIC Science & Technology

2011-12-01

was benchmarked against modern aerodynamic prediction programs like ANSYS CFX and Aero-Prediction 09 (AP09). Next, a comparison was made between two...types of angle of attack generation methods in ANSYS CFX . The research then focused on controlled tilting of the projectile’s nose to investigate the...resulting aerodynamic effects. ANSYS CFX was found to provide better agreement with the experimental data than AP09. 14. SUBJECT

Challenges of Aircraft Design Integration

DTIC Science & Technology

2003-03-01

predicted by the conceptual stick model and the full FEM of the Challenger wing without winglets . Advanced aerodynamic wing design methods To design wings...Piperni, E. Laurendeau Advanced Aerodynamics Bombardier Aerospace 400 CMte Vertu Road Dorval, Quebec, Canada, H4S 1Y9 Fassi.Kafyeke @notes.canadair.ca Tel...514) 855-7186 Abstract The design of a modern airplane brings together many disciplines: structures, aerodynamics , controls, systems, propulsion

Aerodynamic interference effects on tilting proprotor aircraft. [using the Green function method

NASA Technical Reports Server (NTRS)

Soohoo, P.; Morino, L.; Noll, R. B.; Ham, N. D.

1977-01-01

The Green's function method was used to study tilting proprotor aircraft aerodynamics with particular application to the problem of the mutual interference of the wing-fuselage-tail-rotor wake configuration. While the formulation is valid for fully unsteady rotor aerodynamics, attention was directed to steady state aerodynamics, which was achieved by replacing the rotor with the actuator disk approximation. The use of an actuator disk analysis introduced a mathematical singularity into the formulation; this problem was studied and resolved. The pressure distribution, lift, and pitching moment were obtained for an XV-15 wing-fuselage-tail rotor configuration at various flight conditions. For the flight configurations explored, the effects of the rotor wake interference on the XV-15 tilt rotor aircraft yielded a reduction in the total lift and an increase in the nose-down pitching moment. This method provides an analytical capability that is simple to apply and can be used to investigate fuselage-tail rotor wake interference as well as to explore other rotor design problem areas.

Unified Aeroacoustics Analysis for High Speed Turboprop Aerodynamics and Noise. Volume 1; Development of Theory for Blade Loading, Wakes, and Noise

NASA Technical Reports Server (NTRS)

Hanson, D. B.

1991-01-01

A unified theory for the aerodynamics and noise of advanced turboprops are presented. Aerodynamic topics include calculation of performance, blade load distribution, and non-uniform wake flow fields. Blade loading can be steady or unsteady due to fixed distortion, counter-rotating wakes, or blade vibration. The aerodynamic theory is based on the pressure potential method and is therefore basically linear. However, nonlinear effects associated with finite axial induction and blade vortex flow are included via approximate methods. Acoustic topics include radiation of noise caused by blade thickness, steady loading (including vortex lift), and unsteady loading. Shielding of the fuselage by its boundary layer and the wing are treated in separate analyses that are compatible but not integrated with the aeroacoustic theory for rotating blades.

Aircraft aerodynamic prediction method for V/STOL transition including flow separation

NASA Technical Reports Server (NTRS)

Gilmer, B. R.; Miner, G. A.; Bristow, D. R.

1983-01-01

A numerical procedure was developed for the aerodynamic force and moment analysis of V/STOL aircraft operating in the transition regime between hover and conventional forward flight. The trajectories, cross sectional area variations, and mass entrainment rates of the jets are calculated by the Adler-Baron Jet-in-Crossflow Program. The inviscid effects of the interaction between the jets and airframe on the aerodynamic properties are determined by use of the MCAIR 3-D Subsonic properties are determined by use of the MCAIR 3-D Subsonic Potential Flow Program, a surface panel method. In addition, the MCAIR 3-D Geometry influence Coefficient Program is used to calculate a matrix of partial derivatives that represent the rate of change of the inviscid aerodynamic properties with respect to arbitrary changes in the effective wing shape.

Calculated Low-Speed Steady and Time-Dependent Aerodynamic Derivatives for Some Airfoils Using a Discrete Vortex Method

NASA Technical Reports Server (NTRS)

Riley, Donald R.

2015-01-01

This paper contains a collection of some results of four individual studies presenting calculated numerical values for airfoil aerodynamic stability derivatives in unseparated inviscid incompressible flow due separately to angle-of-attack, pitch rate, flap deflection, and airfoil camber using a discrete vortex method. Both steady conditions and oscillatory motion were considered. Variables include the number of vortices representing the airfoil, the pitch axis / moment center chordwise location, flap chord to airfoil chord ratio, and circular or parabolic arc camber. Comparisons with some experimental and other theoretical information are included. The calculated aerodynamic numerical results obtained using a limited number of vortices provided in each study compared favorably with thin airfoil theory predictions. Of particular interest are those aerodynamic results calculated herein (such as induced drag) that are not readily available elsewhere.

Pressure-Sensitive Paint Measurements on Surfaces with Non-Uniform Temperature

NASA Technical Reports Server (NTRS)

Bencic, Timothy J.

1999-01-01

Pressure-sensitive paint (PSP) has become a useful tool to augment conventional pressure taps in measuring the surface pressure distribution of aerodynamic components in wind tunnel testing. While the PSP offers the advantage of a non-intrusive global mapping of the surface pressure, one prominent drawback to the accuracy of this technique is the inherent temperature sensitivity of the coating's luminescent intensity. A typical aerodynamic surface PSP test has relied on the coated surface to be both spatially and temporally isothermal, along with conventional instrumentation for an in situ calibration to generate the highest accuracy pressure mappings. In some tests however, spatial and temporal thermal gradients are generated by the nature of the test as in a blowing jet impinging on a surface. In these cases, the temperature variations on the painted surface must be accounted for in order to yield high accuracy and reliable data. A new temperature correction technique was developed at NASA Lewis to collapse a "family" of PSP calibration curves to a single intensity ratio versus pressure curve. This correction allows a streamlined procedure to be followed whether or not temperature information is used in the data reduction of the PSP. This paper explores the use of conventional instrumentation such as thermocouples and pressure taps along with temperature-sensitive paint (TSP) to correct for the thermal gradients that exist in aeropropulsion PSP tests. Temperature corrected PSP measurements for both a supersonic mixer ejector and jet cavity interaction tests are presented.

Two-Dimensional High-Lift Aerodynamic Optimization Using Neural Networks

NASA Technical Reports Server (NTRS)

Greenman, Roxana M.

1998-01-01

The high-lift performance of a multi-element airfoil was optimized by using neural-net predictions that were trained using a computational data set. The numerical data was generated using a two-dimensional, incompressible, Navier-Stokes algorithm with the Spalart-Allmaras turbulence model. Because it is difficult to predict maximum lift for high-lift systems, an empirically-based maximum lift criteria was used in this study to determine both the maximum lift and the angle at which it occurs. The 'pressure difference rule,' which states that the maximum lift condition corresponds to a certain pressure difference between the peak suction pressure and the pressure at the trailing edge of the element, was applied and verified with experimental observations for this configuration. Multiple input, single output networks were trained using the NASA Ames variation of the Levenberg-Marquardt algorithm for each of the aerodynamic coefficients (lift, drag and moment). The artificial neural networks were integrated with a gradient-based optimizer. Using independent numerical simulations and experimental data for this high-lift configuration, it was shown that this design process successfully optimized flap deflection, gap, overlap, and angle of attack to maximize lift. Once the neural nets were trained and integrated with the optimizer, minimal additional computer resources were required to perform optimization runs with different initial conditions and parameters. Applying the neural networks within the high-lift rigging optimization process reduced the amount of computational time and resources by 44% compared with traditional gradient-based optimization procedures for multiple optimization runs.

Spacecraft stability and control

NASA Technical Reports Server (NTRS)

Barret, Chris

1992-01-01

The Earth's first artificial satellite, Sputnik 1, slowly tumbled in orbit. The first U.S. satellite, Explorer 1, also tumbled out of control. Today, satellite stability and control has become a higher priority. For a satellite design that is to have a life expectancy of 14 years, appropriate spacecraft flight control systems will be reviewed, stability requirements investigated, and an appropriate flight control system recommended in order to see the design process. Disturbance torques, including aerodynamic, magnetic, gravity gradient, solar, micrometeorite, debris, collision, and internal torques, will be assessed to quantify the disturbance environment so that the required compensating torques can be determined. The control torques, including passive versus active, momentum control, bias momentum, spin stabilization, dual spin, gravity gradient, magnetic, reaction wheels, control moment gyros, inertia augmentation techniques, three-axis control, and reaction control systems (RCSs), will be considered. Conditions for stability will also be considered.

The finite element method in low speed aerodynamics

NASA Technical Reports Server (NTRS)

Baker, A. J.; Manhardt, P. D.

1975-01-01

The finite element procedure is shown to be of significant impact in design of the 'computational wind tunnel' for low speed aerodynamics. The uniformity of the mathematical differential equation description, for viscous and/or inviscid, multi-dimensional subsonic flows about practical aerodynamic system configurations, is utilized to establish the general form of the finite element algorithm. Numerical results for inviscid flow analysis, as well as viscous boundary layer, parabolic, and full Navier Stokes flow descriptions verify the capabilities and overall versatility of the fundamental algorithm for aerodynamics. The proven mathematical basis, coupled with the distinct user-orientation features of the computer program embodiment, indicate near-term evolution of a highly useful analytical design tool to support computational configuration studies in low speed aerodynamics.

On-orbit free molecular flow aerodynamic characteristics of a proposal space operations center configuration

NASA Technical Reports Server (NTRS)

Romere, P. O.

1982-01-01

A proposed configuration for a Space Operations Center is presented in its eight stages of buildup. The on orbit aerodynamic force and moment characteristics were calculated for each stage based upon free molecular flow theory. Calculation of the aerodynamic characteristics was accomplished through the use of an orbital aerodynamic computer program, and the computation method is described with respect to the free molecular theory used. The aerodynamic characteristics are presented in tabulated form for each buildup stage at angles of attack from 0 to 360 degrees and roll angles from -60 to +60 degrees. The reference altitude is 490 kilometers, however, the data should be applicable for altitudes below 490 kilometers down to approximately 185 kilometers.

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.

Influence of measuring algorithm on shape accuracy in the compensating turning of high gradient thin-wall parts

NASA Astrophysics Data System (ADS)

Wang, Tao; Wang, Guilin; Zhu, Dengchao; Li, Shengyi

2015-02-01

In order to meet the requirement of aerodynamics, the infrared domes or windows with conformal and thin-wall structure becomes the development trend of high-speed aircrafts in the future. But these parts usually have low stiffness, the cutting force will change along with the axial position, and it is very difficult to meet the requirement of shape accuracy by single machining. Therefore, on-machine measurement and compensating turning are used to control the shape errors caused by the fluctuation of cutting force and the change of stiffness. In this paper, on the basis of ultra precision diamond lathe, a contact measuring system with five DOFs is developed to achieve on-machine measurement of conformal thin-wall parts with high accuracy. According to high gradient surface, the optimizing algorithm is designed on the distribution of measuring points by using the data screening method. The influence rule of sampling frequency is analyzed on measuring errors, the best sampling frequency is found out based on planning algorithm, the effect of environmental factors and the fitting errors are controlled within lower range, and the measuring accuracy of conformal dome is greatly improved in the process of on-machine measurement. According to MgF2 conformal dome with high gradient, the compensating turning is implemented by using the designed on-machine measuring algorithm. The shape error is less than PV 0.8μm, greatly superior compared with PV 3μm before compensating turning, which verifies the correctness of measuring algorithm.

Validation of the virtual elevation field test method when assessing the aerodynamics of para-cyclists with a uni-lateral trans-tibial amputation.

PubMed

Dyer, Bryce; Disley, B Xavier

2018-02-01

Lower-limb amputees typically require some form of prosthetic limb to ride a bicycle for recreation or when competing. At elite-level racing speeds, aerodynamic drag can represent the majority of the resistance acting against a cyclists' forward motion. As a result, the reduction of such resistance is beneficial to an amputee whereby the form and function of the prosthetic limb can be optimized through engineering. To measure the performance of such limbs, field testing provides a cost-effective and context-specific method of aerodynamic drag measurement. However, few methods have been formally validated and none have been applied to amputees with lower-limb amputations. In this paper, an elite level para-cyclist wore two different prosthetic limb designs and had their total aerodynamic drag of a wind tunnel reference method statistically correlated against a velodrome-based virtual elevation field test method. The calculated coefficient of variation was in the range of 0.7-0.9% for the wind tunnel method and 2-3% for the virtual elevation method. A 0.03 m 2 difference was identified in the absolute values recorded between the two methods. Ultimately, both methods exhibited high levels of precision, yet relative results to each other. The virtual elevation method is proposed as a suitable technique to assess the aerodynamic drag of amputee para-cyclists. Implications for rehabilitation This assessment method will provide practitioners a reliable means of assessing the impact of changes made to prosthetics design for cyclists with limb absence. The proposed method offers a low cost and geographically accessible solution compared to others proposed in the past. This assessment method has significant potential for impact among prosthetic limb users looking to improve their cycling performance whereas previous attention in this field has been extremely limited.

An experimental investigation of the aerodynamics and cooling of a horizontally-opposed air-cooled aircraft engine installation

NASA Technical Reports Server (NTRS)

Miley, S. J.; Cross, E. J., Jr.; Owens, J. K.; Lawrence, D. L.

1981-01-01

A flight-test based research program was performed to investigate the aerodynamics and cooling of a horizontally-opposed engine installation. Specific areas investigated were the internal aerodynamics and cooling mechanics of the installation, inlet aerodynamics, and exit aerodynamics. The applicable theory and current state of the art are discussed for each area. Flight-test and ground-test techniques for the development of the cooling installation and the solution of cooling problems are presented. The results show that much of the internal aerodynamics and cooling technology developed for radial engines are applicable to horizontally opposed engines. Correlation is established between engine manufacturer's cooling design data and flight measurements of the particular installation. Also, a flight-test method for the development of cooling requirements in terms of easily measurable parameters is presented. The impact of inlet and exit design on cooling and cooling drag is shown to be of major significance.

A Synthesis of Hybrid RANS/LES CFD Results for F-16XL Aircraft Aerodynamics

NASA Technical Reports Server (NTRS)

Luckring, James M.; Park, Michael A.; Hitzel, Stephan M.; Jirasek, Adam; Lofthouse, Andrew J.; Morton, Scott A.; McDaniel, David R.; Rizzi, Arthur M.

2015-01-01

A synthesis is presented of recent numerical predictions for the F-16XL aircraft flow fields and aerodynamics. The computational results were all performed with hybrid RANS/LES formulations, with an emphasis on unsteady flows and subsequent aerodynamics, and results from five computational methods are included. The work was focused on one particular low-speed, high angle-of-attack flight test condition, and comparisons against flight-test data are included. This work represents the third coordinated effort using the F-16XL aircraft, and a unique flight-test data set, to advance our knowledge of slender airframe aerodynamics as well as our capability for predicting these aerodynamics with advanced CFD formulations. The prior efforts were identified as Cranked Arrow Wing Aerodynamics Project International, with the acronyms CAWAPI and CAWAPI-2. All information in this paper is in the public domain.

Application of the scalar and vector potentials to the aerodynamics of jets

NASA Technical Reports Server (NTRS)

Russell, H. L.; Skifstad, J. G.

1973-01-01

The applicability of a method based on the Stokes potentials (vector and scalar potentials) to computations associated with the aerodynamics of jets was examined. The aerodynamic field near the nozzle could be represented and that the influence of a nonuniform velocity profile at the nozzle exit plane could be determined. Also computations were made for an axisymmetric jet exhausting into a quiescient atmosphere. The velocity at the axis of the jet, and the location of the half-velocity points along the jet yield accurate aerodynamic field computations. Inconsistencies among the different theoretical characterizations of jet flowfields are shown.

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.

High speed propeller acoustics and aerodynamics - A boundary element approach

NASA Technical Reports Server (NTRS)

Farassat, F.; Myers, M. K.; Dunn, M. H.

1989-01-01

The Boundary Element Method (BEM) is applied in this paper to the problems of acoustics and aerodynamics of high speed propellers. The underlying theory is described based on the linearized Ffowcs Williams-Hawkings equation. The surface pressure on the blade is assumed unknown in the aerodynamic problem. It is obtained by solving a singular integral equation. The acoustic problem is then solved by moving the field point inside the fluid medium and evaluating some surface and line integrals. Thus the BEM provides a powerful technique in calculation of high speed propeller aerodynamics and acoustics.

Morphing Wings: A Study Using High-Fidelity Aerodynamic Shape Optimization

NASA Astrophysics Data System (ADS)

Curiale, Nathanael J.

With the aviation industry under pressure to reduce fuel consumption, morphing wings have the capacity to improve aircraft performance, thereby making a significant contribution to reversing climate change. Through high-fidelity aerodynamic shape optimization, various forms of morphing wings are assessed for a hypothetical regional-class aircraft. The framework used solves the Reynolds-averaged Navier-Stokes equations and utilizes a gradient-based optimization algorithm. Baseline geometries are developed through multipoint optimization, where the average drag coefficient is minimized over a range of flight conditions with additional dive constraints. Morphing optimizations are then performed, beginning with these baseline shapes. Five distinct types of morphing are investigated and compared. Overall, a theoretical fully adaptable wing produces roughly a 2% improvement in average performance, whereas trailing-edge morphing with a 27-point multipoint baseline results in just over a 1% improvement in average performance. Trailing-edge morphing proves to be more beneficial than leading-edge morphing, upper-surface morphing, and a conventional flap.

Active two-phase cooling of an IR window for a hypersonic interceptor

NASA Astrophysics Data System (ADS)

Burzlaff, B. H.; Chivian, Jay S.; Cotten, W. D.; Hemphill, R. B.; Huhlein, Michael A.

1993-06-01

A novel actively cooled window for an IR sensor on a hypersonic interceptor is envisioned which achieves an IR window with high transmittance, low emittance, and low image distortion under high aerodynamic heat flux. The cooling concept employs two-phase convective boiling of liquid ammonia. Coolant is confined to narrow, parallel channels within the window to minimize obscuration of the aperture. The high latent heat of vaporization of ammonia minimizes coolant mass-flow requirements. Low boiling temperatures at projected operating pressures promote high thermal conductivity and low emissivity in the window. The concept was tested with thermal measurements on sub-mm width coolant channels in Si. High values for heat transfer coefficient and critical heat flux were obtained. Thermal gradients within the window can be controlled by the coolant channel configuration. Design options are investigated by predicting the effect of aerodynamic heat flux on the image produced by an IR sensor with a cooled window. Ammonia-cooled IR windows will function in the anticipated aerothermal environment.

Multidisciplinary Aerodynamic-Structural Shape Optimization Using Deformation (MASSOUD)

NASA Technical Reports Server (NTRS)

Samareh, Jamshid A.

2000-01-01

This paper presents a multidisciplinary shape parameterization approach. The approach consists of two basic concepts: (1) parameterizing the shape perturbations rather than the geometry itself and (2) performing the shape deformation by means of the soft object animation algorithms used in computer graphics. Because the formulation presented in this paper is independent of grid topology, we can treat computational fluid dynamics and finite element grids in the same manner. The proposed approach is simple, compact, and efficient. Also, the analytical sensitivity derivatives are easily computed for use in a gradient-based optimization. This algorithm is suitable for low-fidelity (e.g., linear aerodynamics and equivalent laminate plate structures) and high-fidelity (e.g., nonlinear computational fluid dynamics and detailed finite element modeling) analysis tools. This paper contains the implementation details of parameterizing for planform, twist, dihedral, thickness, camber, and free-form surface. Results are presented for a multidisciplinary application consisting of nonlinear computational fluid dynamics, detailed computational structural mechanics, and a simple performance module.

Multidisciplinary Aerodynamic-Structural Shape Optimization Using Deformation (MASSOUD)

NASA Technical Reports Server (NTRS)

Samareh, Jamshid A.

2000-01-01

This paper presents a multidisciplinary shape parameterization approach. The approach consists of two basic concepts: (1) parameterizing the shape perturbations rather than the geometry itself and (2) performing the shape deformation by means of the soft object animation algorithms used in computer graphics. Because the formulation presented in this paper is independent of grid topology, we can treat computational fluid dynamics and finite element grids in a similar manner. The proposed approach is simple, compact, and efficient. Also, the analytical sensitivity derivatives are easily computed for use in a gradient-based optimization. This algorithm is suitable for low-fidelity (e.g., linear aerodynamics and equivalent laminated plate structures) and high-fidelity (e.g., nonlinear computational fluid dynamics and detailed finite element modeling analysis tools. This paper contains the implementation details of parameterizing for planform, twist, dihedral, thickness, camber, and free-form surface. Results are presented for a multidisciplinary design optimization application consisting of nonlinear computational fluid dynamics, detailed computational structural mechanics, and a simple performance module.

Determination of Orbiter and Carrier Aerodynamic Coefficients from Load Cell Measurements

NASA Technical Reports Server (NTRS)

Glenn, G. M.

1976-01-01

A method of determining orbiter and carrier total aerodynamic coefficients from load cell measurements is required to support the inert and the captive active flights of the ALT program. A set of equations expressing the orbiter and carrier total aerodynamic coefficients in terms of the load cell measurements, the sensed dynamics of the Boeing 747 (carrier) aircraft, and the relative geometry of the orbiter/carrier is derived.

Numerical study on aerodynamic damping of floating vertical axis wind turbines

NASA Astrophysics Data System (ADS)

Cheng, Zhengshun; Aagaard Madsen, Helge; Gao, Zhen; Moan, Torgeir

2016-09-01

Harvesting offshore wind energy resources using floating vertical axis wind turbines (VAWTs) has attracted an increasing interest in recent years. Due to its potential impact on fatigue damage, the aerodynamic damping should be considered in the preliminary design of a floating VAWT based on the frequency domain method. However, currently the study on aerodynamic damping of floating VAWTs is very limited. Due to the essential difference in aerodynamic load characteristics, the aerodynamic damping of a floating VAWT could be different from that of a floating horizontal axis wind turbine (HAWT). In this study, the aerodynamic damping of floating VAWTs was studied in a fully coupled manner, and its influential factors and its effects on the motions, especially the pitch motion, were demonstrated. Three straight-bladed floating VAWTs with identical solidity and with a blade number varying from two to four were considered. The aerodynamic damping under steady and turbulent wind conditions were estimated using fully coupled aero-hydro-servo-elastic time domain simulations. It is found that the aerodynamic damping ratio of the considered floating VAWTs ranges from 1.8% to 5.3%. Moreover, the aerodynamic damping is almost independent of the rotor azimuth angle, and is to some extent sensitive to the blade number.

Methodology for sensitivity analysis, approximate analysis, and design optimization in CFD for multidisciplinary applications. [computational fluid dynamics

NASA Technical Reports Server (NTRS)

Taylor, Arthur C., III; Hou, Gene W.

1992-01-01

Fundamental equations of aerodynamic sensitivity analysis and approximate analysis for the two dimensional thin layer Navier-Stokes equations are reviewed, and special boundary condition considerations necessary to apply these equations to isolated lifting airfoils on 'C' and 'O' meshes are discussed in detail. An efficient strategy which is based on the finite element method and an elastic membrane representation of the computational domain is successfully tested, which circumvents the costly 'brute force' method of obtaining grid sensitivity derivatives, and is also useful in mesh regeneration. The issue of turbulence modeling is addressed in a preliminary study. Aerodynamic shape sensitivity derivatives are efficiently calculated, and their accuracy is validated on two viscous test problems, including: (1) internal flow through a double throat nozzle, and (2) external flow over a NACA 4-digit airfoil. An automated aerodynamic design optimization strategy is outlined which includes the use of a design optimization program, an aerodynamic flow analysis code, an aerodynamic sensitivity and approximate analysis code, and a mesh regeneration and grid sensitivity analysis code. Application of the optimization methodology to the two test problems in each case resulted in a new design having a significantly improved performance in the aerodynamic response of interest.

A Study of Wake Development and Structure in Constant Pressure Gradients

NASA Technical Reports Server (NTRS)

Thomas, Flint O.; Nelson, R. C.; Liu, Xiaofeng

2000-01-01

Motivated by the application to high-lift aerodynamics for commercial transport aircraft, a systematic investigation into the response of symmetric/asymmetric planar turbulent wake development to constant adverse, zero, and favorable pressure gradients has been conducted. The experiments are performed at a Reynolds number of 2.4 million based on the chord of the wake generator. A unique feature of this wake study is that the pressure gradients imposed on the wake flow field are held constant. The experimental measurements involve both conventional LDV and hot wire flow field surveys of mean and turbulent quantities including the turbulent kinetic energy budget. In addition, similarity analysis and numerical simulation have also been conducted for this wake study. A focus of the research has been to isolate the effects of both pressure gradient and initial wake asymmetry on the wake development. Experimental results reveal that the pressure gradient has a tremendous influence on the wake development, despite the relatively modest pressure gradients imposed. For a given pressure gradient, the development of an initially asymmetric wake is different from the initially symmetric wake. An explicit similarity solution for the shape parameters of the symmetric wake is obtained and agrees with the experimental results. The turbulent kinetic energy budget measurements of the symmetric wake demonstrate that except for the convection term, the imposed pressure gradient does not change the fundamental flow physics of turbulent kinetic energy transport. Based on the turbulent kinetic energy budget measurements, an approach to correct the bias error associated with the notoriously difficult dissipation estimate is proposed and validated through the comparison of the experimental estimate with a direct numerical simulation result.

Adaptive Aft Signature Shaping of a Low-Boom Supersonic Aircraft Using Off-Body Pressures

NASA Technical Reports Server (NTRS)

Ordaz, Irian; Li, Wu

2012-01-01

The design and optimization of a low-boom supersonic aircraft using the state-of-the- art o -body aerodynamics and sonic boom analysis has long been a challenging problem. The focus of this paper is to demonstrate an e ective geometry parameterization scheme and a numerical optimization approach for the aft shaping of a low-boom supersonic aircraft using o -body pressure calculations. A gradient-based numerical optimization algorithm that models the objective and constraints as response surface equations is used to drive the aft ground signature toward a ramp shape. The design objective is the minimization of the variation between the ground signature and the target signature subject to several geometric and signature constraints. The target signature is computed by using a least-squares regression of the aft portion of the ground signature. The parameterization and the deformation of the geometry is performed with a NASA in- house shaping tool. The optimization algorithm uses the shaping tool to drive the geometric deformation of a horizontal tail with a parameterization scheme that consists of seven camber design variables and an additional design variable that describes the spanwise location of the midspan section. The demonstration cases show that numerical optimization using the state-of-the-art o -body aerodynamic calculations is not only feasible and repeatable but also allows the exploration of complex design spaces for which a knowledge-based design method becomes less effective.

Vortex lattice prediction of subsonic aerodynamics of hypersonic vehicle concepts

NASA Technical Reports Server (NTRS)

Pittman, J. L.; Dillon, J. L.

1977-01-01

The vortex lattice method introduced by Lamar and Gloss (1975) was applied to the prediction of subsonic aerodynamic characteristics of hypersonic body-wing configurations. The reliability of the method was assessed through comparison of the calculated and observed aerodynamic performances of two National Hypersonic Flight Research Facility craft at Mach 0.2. The investigation indicated that a vortex lattice model involving 120 or more panel elements can give good results for the lift and induced drag coefficients of the craft, as well as for the pitching moment at angles of attack below 10 to 15 deg. Automated processes for calculating the local slopes of mean-camber surfaces may also render the method suitable for use in preliminary design phases.

Aerodynamic parameter studies and sensitivity analysis for rotor blades in axial flight

NASA Technical Reports Server (NTRS)

Chiu, Y. Danny; Peters, David A.

1991-01-01

The analytical capability is offered for aerodynamic parametric studies and sensitivity analyses of rotary wings in axial flight by using a 3-D undistorted wake model in curved lifting line theory. The governing equations are solved by both the Multhopp Interpolation technique and the Vortex Lattice method. The singularity from the bound vortices is eliminated through the Hadamard's finite part concept. Good numerical agreement between both analytical methods and finite differences methods are found. Parametric studies were made to assess the effects of several shape variables on aerodynamic loads. It is found, e.g., that a rotor blade with out-of-plane and inplane curvature can theoretically increase lift in the inboard and outboard regions respectively without introducing an additional induced drag.

Methods of the aerodynamical experiments with simulation of massflow-traction ratio of the power unit

NASA Astrophysics Data System (ADS)

Lokotko, A. V.

2016-10-01

Modeling massflow-traction characteristics of the power unit (PU) may be of interest in the study of aerodynamic characteristics (ADC) aircraft models with full dynamic likeness, and in the study of the effect of interference PU. These studies require the use of a number of processing methods. These include: 1) The method of delivery of the high-pressure body of jets model engines on the sensitive part of the aerodynamic balance. 2) The method of estimate accuracy and reliability of measurement thrust generated by the jet device. 3) The method of implementation of the simulator SU in modeling the external contours of the nacelle, and the conditions at the inlet and outlet. 4) The method of determining the traction simulator PU. 5) The method of determining the interference effect from the work of power unit on the ADC of model. 6) The method of producing hot jets of jet engines. The paper examines implemented in ITAM methodology applied to testing in a supersonic wind tunnel T-313.

Aerodynamic Noise Generated by Shinkansen Cars

NASA Astrophysics Data System (ADS)

KITAGAWA, T.; NAGAKURA, K.

2000-03-01

The noise value (A -weighted sound pressure level, SLOW) generated by Shinkansen trains, now running at 220-300 km/h, should be less than 75 dB(A) at the trackside. Shinkansen noise, such as rolling noise, concrete support structure noise, and aerodynamic noise are generated by various parts of Shinkansen trains. Among these aerodynamic noise is important because it is the major contribution to the noise generated by the coaches running at high speed. In order to reduce the aerodynamic noise, a number of improvements to coaches have been made. As a result, the aerodynamic noise has been reduced, but it still remains significant. In addition, some aerodynamic noise generated from the lower parts of cars remains. In order to investigate the contributions of these noises, a method of analyzing Shinkansen noise has been developed and applied to the measured data of Shinkansen noise at speeds between 120 and 315 km/h. As a result, the following conclusions have been drawn: (1) Aerodynamic noise generated from the upper parts of cars was reduced considerably by smoothing car surfaces. (2) Aerodynamic noise generated from the lower parts of cars has a major influence upon the wayside noise.

Simultaneous Aerodynamic Analysis and Design Optimization (SAADO) for a 3-D Flexible Wing

NASA Technical Reports Server (NTRS)

Gumbert, Clyde R.; Hou, Gene J.-W.

2001-01-01

The formulation and implementation of an optimization method called Simultaneous Aerodynamic Analysis and Design Optimization (SAADO) are extended from single discipline analysis (aerodynamics only) to multidisciplinary analysis - in this case, static aero-structural analysis - and applied to a simple 3-D wing problem. The method aims to reduce the computational expense incurred in performing shape optimization using state-of-the-art Computational Fluid Dynamics (CFD) flow analysis, Finite Element Method (FEM) structural analysis and sensitivity analysis tools. Results for this small problem show that the method reaches the same local optimum as conventional optimization. However, unlike its application to the win,, (single discipline analysis), the method. as I implemented here, may not show significant reduction in the computational cost. Similar reductions were seen in the two-design-variable (DV) problem results but not in the 8-DV results given here.

Analysis of iced wings

NASA Technical Reports Server (NTRS)

Cebeci, T.; Chen, H. H.; Kaups, K.; Schimke, S.; Shin, J.

1992-01-01

A method for computing ice shapes along the leading edge of a wing and a method for predicting its aerodynamic performance degradation due to icing is described. Ice shapes are computed using an extension of the LEWICE code which was developed for airfoils. The aerodynamic properties of the iced wing are determined with an interactive scheme in which the solutions of the inviscid flow equations are obtained from a panel method and the solutions of the viscous flow equations are obtained from an inverse three-dimensional finite-difference boundary-layer method. A new interaction law is used to couple the inviscid and viscous flow solutions. The application of the LEWICE wing code to the calculation of ice shapes on a MS-317 swept wing shows good agreement with measurements. The interactive boundary-layer method is applied to a tapered ice wing in order to study the effect of icing on the aerodynamic properties of the wing at several angles of attack.

New approaches to optimization in aerospace conceptual design

NASA Technical Reports Server (NTRS)

Gage, Peter J.

1995-01-01

Aerospace design can be viewed as an optimization process, but conceptual studies are rarely performed using formal search algorithms. Three issues that restrict the success of automatic search are identified in this work. New approaches are introduced to address the integration of analyses and optimizers, to avoid the need for accurate gradient information and a smooth search space (required for calculus-based optimization), and to remove the restrictions imposed by fixed complexity problem formulations. (1) Optimization should be performed in a flexible environment. A quasi-procedural architecture is used to conveniently link analysis modules and automatically coordinate their execution. It efficiently controls a large-scale design tasks. (2) Genetic algorithms provide a search method for discontinuous or noisy domains. The utility of genetic optimization is demonstrated here, but parameter encodings and constraint-handling schemes must be carefully chosen to avoid premature convergence to suboptimal designs. The relationship between genetic and calculus-based methods is explored. (3) A variable-complexity genetic algorithm is created to permit flexible parameterization, so that the level of description can change during optimization. This new optimizer automatically discovers novel designs in structural and aerodynamic tasks.

Variable Camber Continuous Aerodynamic Control Surfaces and Methods for Active Wing Shaping Control

NASA Technical Reports Server (NTRS)

Nguyen, Nhan T. (Inventor)

2016-01-01

An aerodynamic control apparatus for an air vehicle improves various aerodynamic performance metrics by employing multiple spanwise flap segments that jointly form a continuous or a piecewise continuous trailing edge to minimize drag induced by lift or vortices. At least one of the multiple spanwise flap segments includes a variable camber flap subsystem having multiple chordwise flap segments that may be independently actuated. Some embodiments also employ a continuous leading edge slat system that includes multiple spanwise slat segments, each of which has one or more chordwise slat segment. A method and an apparatus for implementing active control of a wing shape are also described and include the determination of desired lift distribution to determine the improved aerodynamic deflection of the wings. Flap deflections are determined and control signals are generated to actively control the wing shape to approximate the desired deflection.

Modeling of aerodynamic heat flux and thermoelastic behavior of nose caps of hypersonic vehicles

NASA Astrophysics Data System (ADS)

Persova, Marina G.; Soloveichik, Yury G.; Belov, Vasiliy K.; Kiselev, Dmitry S.; Vagin, Denis V.; Domnikov, Petr A.; Patrushev, Ilya I.; Kurskiy, Denis N.

2017-07-01

In this paper, the problem of numerical modeling of thermoelastic behavior of nose caps of hypersonic vehicles at different angles of attack is considered. 3D finite element modeling is performed by solving the coupled heat and elastic problems taking into account thermal and mechanical properties variations with temperature. A special method for calculating the aerodynamic heat flux entering the nose cap from its surface is proposed. This method is characterized by very low computational costs and allows calculating the aerodynamic heat flux at different values of the Mach number and angles of attack which may vary during the aerodynamic heating. The numerical results obtained by the proposed approach are compared with the numerical results and experimental data obtained by other authors. The developed approach has been used for studying the impact of the angle of attack on the thermoelastic behavior of nose caps main components.

Extension of HCDstruct for Transonic Aeroservoelastic Analysis of Unconventional Aircraft Concepts

NASA Technical Reports Server (NTRS)

Quinlan, Jesse R.; Gern, Frank H.

2017-01-01

A substantial effort has been made to implement an enhanced aerodynamic modeling capability in the Higher-fidelity Conceptual Design and structural optimization tool. This additional capability is needed for a rapid, physics-based method of modeling advanced aircraft concepts at risk of structural failure due to dynamic aeroelastic instabilities. To adequately predict these instabilities, in particular for transonic applications, a generalized aerodynamic matching algorithm was implemented to correct the doublet-lattice model available in Nastran using solution data from a priori computational fluid dynamics anal- ysis. This new capability is demonstrated for two tube-and-wing aircraft configurations, including a Boeing 737-200 for implementation validation and the NASA D8 as a first use case. Results validate the current implementation of the aerodynamic matching utility and demonstrate the importance of using such a method for aircraft configurations featuring fuselage-wing aerodynamic interaction.

A new method for aerodynamic test of high altitude propellers

NASA Astrophysics Data System (ADS)

Gong, Xiying; Zhang, Lin

A ground test system is designed for aerodynamic performance tests of high altitude propellers. The system is consisted of stable power supply, servo motors, two-component balance constructed by tension-compression sensors, ultrasonic anemometer, data acquisition module. It is loaded on a truck to simulate propellers' wind-tunnel test for different wind velocities at low density circumstance. The graphical programming language LABVIEW for developing virtual instrument is used to realize the test system control and data acquisition. Aerodynamic performance test of a propeller with 6.8 m diameter was completed by using this system. The results verify the feasibility of the ground test method.

Aerodynamic and structural studies of joined-wing aircraft

NASA Technical Reports Server (NTRS)

Kroo, Ilan; Smith, Stephen; Gallman, John

1991-01-01

A method for rapidly evaluating the structural and aerodynamic characteristics of joined-wing aircraft was developed and used to study the fundamental advantages attributed to this concept. The technique involves a rapid turnaround aerodynamic analysis method for computing minimum trimmed drag combined with a simple structural optimization. A variety of joined-wing designs are compared on the basis of trimmed drag, structural weight, and, finally, trimmed drag with fixed structural weight. The range of joined-wing design parameters resulting in best cruise performance is identified. Structural weight savings and net drag reductions are predicted for certain joined-wing configurations compared with conventional cantilever-wing configurations.

An experimental study of an airfoil with a bio-inspired leading edge device at high angles of attack

NASA Astrophysics Data System (ADS)

Mandadzhiev, Boris A.; Lynch, Michael K.; Chamorro, Leonardo P.; Wissa, Aimy A.

2017-09-01

Robust and predictable aerodynamic performance of unmanned aerial vehicles at the limits of their design envelope is critical for safety and mission adaptability. Deployable aerodynamic surfaces from the wing leading or trailing edges are often used to extend the aerodynamic envelope (e.g. slats and flaps). Birds have also evolved feathers at the leading edge (LE) of their wings, known as the alula, which enables them to perform high angles of attack maneuvers. In this study, a series of wind tunnel experiments are performed to quantify the effect of various deployment parameters of an alula-like LE device on the aerodynamic performance of a cambered airfoil (S1223) at stall and post stall conditions. The alula relative angle of attack, measured from the mean chord of the airfoil, is varied to modulate tip-vortex strength, while the alula deflection angle is varied to modulate the distance between the tip vortex and the wing surface. Integrated lift force measurements were collected at various alula-inspired device configurations. The effect of the alula-inspired device on the boundary layer velocity profile and turbulence intensity were investigated through hot-wire anemometer measurements. Results show that as alula deflection angle increases, the lift coefficient also increase especially at lower alula relative angles of attack. Moreover, at post stall wing angles of attack, the wake velocity deficit is reduced in the presence of alula device, confirming the mitigation of the wing adverse pressure gradient. The results are in strong agreement with measurements taken on bird wings showing delayed flow reversal and extended range of operational angles of attack. An engineered alula-inspired device has the potential to improve mission adaptability in small unmanned air vehicles during low Reynolds number flight.

STS-40 orbital acceleration research experiment flight results during a typical sleep period

NASA Technical Reports Server (NTRS)

Blanchard, R. C.; Nicholson, J. Y.; Ritter, J. R.

1992-01-01

The Orbital Acceleration Research Experiment (OARE), an electrostatic accelerometer package with complete on-orbit calibration capabilities, was flown for the first time aboard the Space Shuttle on STS-40. This is also the first time an accelerometer package with nano-g sensitivity and a calibration facility has flown aboard the Space Shuttle. The instrument is designed to measure and record the Space Shuttle aerodynamic acceleration environment from the free molecule flow regime through the rarified flow transition into the hypersonic continuum regime. Because of its sensitivity, the OARE instrument defects aerodynamic behavior of the Space Shuttle while in low-earth orbit. A 2-hour orbital time period on day seven of the mission, when the crew was asleep and other spacecraft activities were at a minimum, was examined. During the flight, a 'trimmed-mean' filter was used to produce high quality, low frequency data which was successfully stored aboard the Space Shuttle in the OARE data storage system. Initial review of the data indicated that, although the expected precision was achieved, some equipment problems occurred resulting in uncertain accuracy. An acceleration model which includes aerodynamic, gravity-gradient, and rotational effects was constructed and compared with flight data. Examination of the model with the flight data shows the instrument to be sensitive to all major expected low frequency acceleration phenomena; however, some erratic instrument bias behavior persists in two axes. In these axes, the OARE data can be made to match a comprehensive atmospheric-aerodynamic model by making bias adjustments and slight linear corrections for drift. The other axis does not exhibit these difficulties and gives good agreement with the acceleration model.

Comparison of Various Supersonic Turbine Tip Designs to Minimize Aerodynamic Loss and Tip Heating

NASA Technical Reports Server (NTRS)

Shyam, Vikram; Ameri, Ali

2012-01-01

The rotor tips of axial turbines experience high heat flux and are the cause of aerodynamic losses due to tip clearance flows, and in the case of supersonic tips, shocks. As stage loadings increase, the flow in the tip gap approaches and exceeds sonic conditions. This introduces effects such as shock-boundary layer interactions and choked flow that are not observed for subsonic tip flows that have been studied extensively in literature. This work simulates the tip clearance flow for a flat tip, a diverging tip gap and several contoured tips to assess the possibility of minimizing tip heat flux while maintaining a constant massflow from the pressure side to the suction side of the rotor, through the tip clearance. The Computational Fluid Dynamics (CFD) code GlennHT was used for the simulations. Due to the strong favorable pressure gradients the simulations assumed laminar conditions in the tip gap. The nominal tip gap width to height ratio for this study is 6.0. The Reynolds number of the flow is 2.4 x 10(exp 5) based on nominal tip width and exit velocity. A wavy wall design was found to reduce heat flux by 5 percent but suffered from an additional 6 percent in aerodynamic loss coefficient. Conventional tip recesses are found to perform far worse than a flat tip due to severe shock heating. Overall, the baseline flat tip was the second best performer. A diverging converging tip gap with a hole was found to be the best choice. Average tip heat flux was reduced by 37 percent and aerodynamic losses were cut by over 6 percent.

Computational and theoretical investigation of Mars's atmospheric impact on the descent module "Exomars-2018" under aerodynamic deceleration

NASA Astrophysics Data System (ADS)

Golomazov, M. M.; Ivankov, A. A.

2016-12-01

Methods for calculating the aerodynamic impact of the Martian atmosphere on the descent module "Exomars-2018" intended for solving the problem of heat protection of the descent module during aerodynamic deceleration are presented. The results of the investigation are also given. The flow field and radiative and convective heat exchange are calculated along the trajectory of the descent module until parachute system activation.

Conjugate problems of transport phenomena under quasi-steady microaccelerations in realistic spaceflight.

PubMed

Polezhaev, V I; Nikitin, S A

2009-04-01

A new model for spatial convective transport processes conjugated with the measured or calculated realistic quasi-steady microaccelerations is presented. Rotation around the mass center, including accelerated rotation, gravity gradient, and aerodynamical drag are taken into account. New results of the effect on mixing and concentration inhomogeneities of the elementary convective processes are presented. The mixing problem in spacecraft enclosures, concentration inhomogeneities due to convection induced by body forces in realistic spaceflight, and the coupling of this kind of convection with thermocapillary convection on the basis of this model are discussed.

Surface temperature effect on subsonic stall.

NASA Technical Reports Server (NTRS)

Macha, J. M.; Norton, D. J.; Young, J. C.

1972-01-01

Results of an analytical and experimental study of boundary layer flow over an aerodynamic surface rejecting heat to a cool environment. This occurs following reentry of a Space Shuttle vehicle. Analytical studies revealed that a surface to freestream temperature ratio, greater than unity tended to destabilize the boundary layer, hastening transition and separation. Therefore, heat transfer accentuated the effect of an adverse pressure gradient. Wind tunnel tests of a 0012-64 NACA airfoil showed that the stall angle was significantly reduced while drag tended to increase for freestream temperature ratios up to 2.2.

Proceedings of Workshop on Atmospheric Density and Aerodynamic Drag Models for Air Force Operations Held at Air Force Geophysics Laboratory on 20-22 October 1987. Volume 1

DTIC Science & Technology

1990-02-13

considered with these production processes in a simple photochemical equilibrium calculation , we are able to determine the contribution each makes to the...Hessian matrix of second derivatives (which is required in the Newton-Raphson procedure) by the vector product of the gradient (VJ) and its transpose...was focused on the altitude region 80-250 Km. Papers were presented in the folowing areas: Air Force requirements , physics of density and drag

NASA Perspective on Requirements for Development of Advanced Methods Predicting Unsteady Aerodynamics and Aeroelasticity

NASA Technical Reports Server (NTRS)

Schuster, David M.

2008-01-01

Over the past three years, the National Aeronautics and Space Administration (NASA) has initiated design, development, and testing of a new human-rated space exploration system under the Constellation Program. Initial designs within the Constellation Program are scheduled to replace the present Space Shuttle, which is slated for retirement within the next three years. The development of vehicles for the Constellation system has encountered several unsteady aerodynamics challenges that have bearing on more traditional unsteady aerodynamic and aeroelastic analysis. This paper focuses on the synergy between the present NASA challenges and the ongoing challenges that have historically been the subject of research and method development. There are specific similarities in the flows required to be analyzed for the space exploration problems and those required for some of the more nonlinear unsteady aerodynamic and aeroelastic problems encountered on aircraft. The aggressive schedule, significant technical challenge, and high-priority status of the exploration system development is forcing engineers to implement existing tools and techniques in a design and application environment that is significantly stretching the capability of their methods. While these methods afford the users with the ability to rapidly turn around designs and analyses, their aggressive implementation comes at a price. The relative immaturity of the techniques for specific flow problems and the inexperience with their broad application to them, particularly on manned spacecraft flight system, has resulted in the implementation of an extensive wind tunnel and flight test program to reduce uncertainty and improve the experience base in the application of these methods. This provides a unique opportunity for unsteady aerodynamics and aeroelastic method developers to test and evaluate new analysis techniques on problems with high potential for acquisition of test and even flight data against which they can be evaluated. However, researchers may be required to alter the geometries typically used in their analyses, the types of flows analyzed, and even the techniques by which computational tools are verified and validated. This paper discusses these issues and provides some perspective on the potential for new and innovative approaches to the development of methods to attack problems in nonlinear unsteady aerodynamics.

Isoprene emission rates and fluxes measured above a Mediterranean oak ( Quercus pubescens) forest

NASA Astrophysics Data System (ADS)

Simon, V.; Dumergues, L.; Bouchou, P.; Torres, L.; Lopez, A.

2005-03-01

The present work, carried out as part of the European fiEld experimentS to COnstrain Models of atmospheric Pollution and Transport of Emissions project (ESCOMPTE), brings a new contribution to the inventory of the main natural hydrocarbons sources that are liable to participate in the production of ozone. The measurement campaign was conducted in Montmeyan, a site close to Marseilles (France), with the aim of quantifying the terpenic emission pattern and the behaviour of Quercus pubescens, an important Mediterranean tree species. Biogenic emissions by Q. pubescens were determined by the enclosure of an intact branch of this tree in a Teflon cuvette. The total monoterpenic emission rates thus recorded were found to reach maximum values ranged between 40 and 350 μg g Dry Weight-1 h -1. Emissions were correlated strongly with leaf temperature and Photosynthetic Active Radiation (PAR). The fluxes were also determined by extrapolating the results of the enclosure method and by using aerodynamic gradient method. They reach around 73 mg m -2 h -1 with the first method and 55 mg m -2 h -1 with the second one. The obtained values fit with a maximal ratio of 2.

High resolution flow field prediction for tail rotor aeroacoustics

NASA Technical Reports Server (NTRS)

Quackenbush, Todd R.; Bliss, Donald B.

1989-01-01

The prediction of tail rotor noise due to the impingement of the main rotor wake poses a significant challenge to current analysis methods in rotorcraft aeroacoustics. This paper describes the development of a new treatment of the tail rotor aerodynamic environment that permits highly accurate resolution of the incident flow field with modest computational effort relative to alternative models. The new approach incorporates an advanced full-span free wake model of the main rotor in a scheme which reconstructs high-resolution flow solutions from preliminary, computationally inexpensive simulations with coarse resolution. The heart of the approach is a novel method for using local velocity correction terms to capture the steep velocity gradients characteristic of the vortex-dominated incident flow. Sample calculations have been undertaken to examine the principal types of interactions between the tail rotor and the main rotor wake and to examine the performance of the new method. The results of these sample problems confirm the success of this approach in capturing the high-resolution flows necessary for analysis of rotor-wake/rotor interactions with dramatically reduced computational cost. Computations of radiated sound are also carried out that explore the role of various portions of the main rotor wake in generating tail rotor noise.

PROP3D: A Program for 3D Euler Unsteady Aerodynamic and Aeroelastic (Flutter and Forced Response) Analysis of Propellers. Version 1.0

NASA Technical Reports Server (NTRS)

Srivastava, R.; Reddy, T. S. R.

1996-01-01

This guide describes the input data required, for steady or unsteady aerodynamic and aeroelastic analysis of propellers and the output files generated, in using PROP3D. The aerodynamic forces are obtained by solving three dimensional unsteady, compressible Euler equations. A normal mode structural analysis is used to obtain the aeroelastic equations, which are solved using either time domain or frequency domain solution method. Sample input and output files are included in this guide for steady aerodynamic analysis of single and counter-rotation propellers, and aeroelastic analysis of single-rotation propeller.

Compendium of NASA Langley reports on hypersonic aerodynamics

NASA Technical Reports Server (NTRS)

Sabo, Frances E.; Cary, Aubrey M.; Lawson, Shirley W.

1987-01-01

Reference is made to papers published by the Langley Research Center in various areas of hypersonic aerodynamics for the period 1950 to 1986. The research work was performed either in-house by the Center staff or by other personnel supported entirely or in part by grants or contracts. Abstracts have been included with the references when available. The references are listed chronologically and are grouped under the following general headings: (1) Aerodynamic Measurements - Single Shapes; (2) Aerodynamic Measurements - Configurations; (3) Aero-Heating; (4) Configuration Studies; (5) Propulsion Integration Experiment; (6) Propulsion Integration - Study; (7) Analysis Methods; (8) Test Techniques; and (9) Airframe Active Cooling Systems.

Finding optimum airfoil shape to get maximum aerodynamic efficiency for a wind turbine

NASA Astrophysics Data System (ADS)

Sogukpinar, Haci; Bozkurt, Ismail

2017-02-01

In this study, aerodynamic performances of S-series wind turbine airfoil of S 825 are investigated to find optimum angle of attack. Aerodynamic performances calculations are carried out by utilization of a Computational Fluid Dynamics (CFD) method withstand finite capacity approximation by using Reynolds-Averaged-Navier Stokes (RANS) theorem. The lift and pressure coefficients, lift to drag ratio of airfoil S 825 are analyzed with SST turbulence model then obtained results crosscheck with wind tunnel data to verify the precision of computational Fluid Dynamics (CFD) approximation. The comparison indicates that SST turbulence model used in this study can predict aerodynamics properties of wind blade.

An Aerodynamic Analysis of a Spinning Missile with Dithering Canards

NASA Technical Reports Server (NTRS)

Meakin, Robert L.; Nygaard, Tor A.

2003-01-01

A generic spinning missile with dithering canards is used to demonstrate the utility of an overset structured grid approach for simulating the aerodynamics of rolling airframe missile systems. The approach is used to generate a modest aerodynamic database for the generic missile. The database is populated with solutions to the Euler and Navier-Stokes equations. It is used to evaluate grid resolution requirements for accurate prediction of instantaneous missile loads and the relative aerodynamic significance of angle-of-attack, canard pitching sequence, viscous effects, and roll-rate effects. A novel analytical method for inter- and extrapolation of database results is also given.

Theory of wing rock

NASA Technical Reports Server (NTRS)

Hsu, C. H.; Lan, C. E.

1984-01-01

A theory is developed for predicting wing rock characteristics. From available data, it can be concluded that wing rock is triggered by flow asymmetries, developed by negative or weakly positive roll damping, and sustained by nonlinear aerodynamic roll damping. A new nonlinear aerodynamic model that includes all essential aerodynamic nonlinearities is developed. The Beecham-Titchener method is applied to obtain approximate analytic solutions for the amplitude and frequency of the limit cycle based on the three degree-of-freedom equations of motion. An iterative scheme is developed to calculate the average aerodynamic derivatives and dynamic characteristics at limit cycle conditions. Good agreement between theoretical and experimental results is obtained.

Method of performing computational aeroelastic analyses

NASA Technical Reports Server (NTRS)

Silva, Walter A. (Inventor)

2011-01-01

Computational aeroelastic analyses typically use a mathematical model for the structural modes of a flexible structure and a nonlinear aerodynamic model that can generate a plurality of unsteady aerodynamic responses based on the structural modes for conditions defining an aerodynamic condition of the flexible structure. In the present invention, a linear state-space model is generated using a single execution of the nonlinear aerodynamic model for all of the structural modes where a family of orthogonal functions is used as the inputs. Then, static and dynamic aeroelastic solutions are generated using computational interaction between the mathematical model and the linear state-space model for a plurality of periodic points in time.

Shape optimization for aerodynamic efficiency and low observability

NASA Technical Reports Server (NTRS)

Vinh, Hoang; Van Dam, C. P.; Dwyer, Harry A.

1993-01-01

Field methods based on the finite-difference approximations of the time-domain Maxwell's equations and the potential-flow equation have been developed to solve the multidisciplinary problem of airfoil shaping for aerodynamic efficiency and low radar cross section (RCS). A parametric study and an optimization study employing the two analysis methods are presented to illustrate their combined capabilities. The parametric study shows that for frontal radar illumination, the RCS of an airfoil is independent of the chordwise location of maximum thickness but depends strongly on the maximum thickness, leading-edge radius, and leadingedge shape. In addition, this study shows that the RCS of an airfoil can be reduced without significant effects on its transonic aerodynamic efficiency by reducing the leading-edge radius and/or modifying the shape of the leading edge. The optimization study involves the minimization of wave drag for a non-lifting, symmetrical airfoil with constraints on the airfoil maximum thickness and monostatic RCS. This optimization study shows that the two analysis methods can be used effectively to design aerodynamically efficient airfoils with certain desired RCS characteristics.

Nonlinear programming extensions to rational function approximations of unsteady aerodynamics

NASA Technical Reports Server (NTRS)

Tiffany, Sherwood H.; Adams, William M., Jr.

1987-01-01

This paper deals with approximating unsteady generalized aerodynamic forces in the equations of motion of a flexible aircraft. Two methods of formulating these approximations are extended to include both the same flexibility in constraining them and the same methodology in optimizing nonlinear parameters as another currently used 'extended least-squares' method. Optimal selection of 'nonlinear' parameters is made in each of the three methods by use of the same nonlinear (nongradient) optimizer. The objective of the nonlinear optimization is to obtain rational approximations to the unsteady aerodynamics whose state-space realization is of lower order than that required when no optimization of the nonlinear terms is performed. The free 'linear' parameters are determined using least-squares matrix techniques on a Lagrange multiplier formulation of an objective function which incorporates selected linear equality constraints. State-space mathematical models resulting from the different approaches are described, and results are presented which show comparative evaluations from application of each of the extended methods to a numerical example. The results obtained for the example problem show a significant (up to 63 percent) reduction in the number of differential equations used to represent the unsteady aerodynamic forces in linear time-invariant equations of motion as compared to a conventional method in which nonlinear terms are not optimized.

Comparison of analytical and experimental steadyand unsteady-pressure distributions at Mach number 0.78 for a high-aspect-ratio supercritical wing model with oscillating control surfaces

NASA Technical Reports Server (NTRS)

Mccain, W. E.

1984-01-01

The unsteady aerodynamic lifting surface theory, the Doublet Lattice method, with experimental steady and unsteady pressure measurements of a high aspect ratio supercritical wing model at a Mach number of 0.78 were compared. The steady pressure data comparisons were made for incremental changes in angle of attack and control surface deflection. The unsteady pressure data comparisons were made at set angle of attack positions with oscillating control surface deflections. Significant viscous and transonic effects in the experimental aerodynamics which cannot be predicted by the Doublet Lattice method are shown. This study should assist development of empirical correction methods that may be applied to improve Doublet Lattice calculations of lifting surface aerodynamics.

Study of flutter related computational procedures for minimum weight structural sizing of advanced aircraft, supplemental data

NASA Technical Reports Server (NTRS)

Oconnell, R. F.; Hassig, H. J.; Radovcich, N. A.

1975-01-01

Computational aspects of (1) flutter optimization (minimization of structural mass subject to specified flutter requirements), (2) methods for solving the flutter equation, and (3) efficient methods for computing generalized aerodynamic force coefficients in the repetitive analysis environment of computer-aided structural design are discussed. Specific areas included: a two-dimensional Regula Falsi approach to solving the generalized flutter equation; method of incremented flutter analysis and its applications; the use of velocity potential influence coefficients in a five-matrix product formulation of the generalized aerodynamic force coefficients; options for computational operations required to generate generalized aerodynamic force coefficients; theoretical considerations related to optimization with one or more flutter constraints; and expressions for derivatives of flutter-related quantities with respect to design variables.

Using the attitude response of aerostable spacecraft to measure thermospheric wind

NASA Astrophysics Data System (ADS)

Virgili-Llop, Josep; Roberts, Peter C. E.; Hao, Zhou

2018-03-01

In situ measurements of the thermospheric wind can be obtained by observing the attitude response of an aerostable spacecraft. In the proposed method, the aerostable spacecraft is left uncontrolled, freely reacting to the aerodynamic torques, and oscillating around its equilibrium attitude. The wind's magnitude and direction is determined by combining the attitude observations with estimates of the other perturbing torques, atmospheric density, and spacecraft's aerodynamic properties. The spatial resolution of the measurements is proportional to the natural frequency of the attitude's oscillation. Spacecraft with high aerodynamic stiffness to inertia ratios operating at low altitudes exhibit higher natural frequencies, making them particularly suited for this method. A one degree-of-freedom case is used to present and illustrate the proposed method as well as to analyze its performance.

A method for the design of transonic flexible wings

NASA Technical Reports Server (NTRS)

Smith, Leigh Ann; Campbell, Richard L.

1990-01-01

Methodology was developed for designing airfoils and wings at transonic speeds which includes a technique that can account for static aeroelastic deflections. This procedure is capable of designing either supercritical or more conventional airfoil sections. Methods for including viscous effects are also illustrated and are shown to give accurate results. The methodology developed is an interactive system containing three major parts. A design module was developed which modifies airfoil sections to achieve a desired pressure distribution. This design module works in conjunction with an aerodynamic analysis module, which for this study is a small perturbation transonic flow code. Additionally, an aeroelastic module is included which determines the wing deformation due to the calculated aerodynamic loads. Because of the modular nature of the method, it can be easily coupled with any aerodynamic analysis code.

International Conference on the Methods of Aerophysical Research 98 "ICMAR 98". Part 3: Proceedings

DTIC Science & Technology

1998-01-01

the study of aerodynamic characteristics and heat transfer in the simplest machine, disk fan, are presented in the paper. FLOW PATTERNS New knowledge of... Aerodynamics of base combustion / Ed. S.N.B.Murthy. - New York: AIAA. 1976. 2. Baev V.K., Golovichev V.I., Tretyakov P.K. Combustion in Supersonic...Theoretical and Applied Mechanics SB RAS, 630090, Novosibirsk, Russia The development of short-duration aerodynamic wind tunnels requires perfection of

AEROX: Computer program for transonic aircraft aerodynamics to high angles of attack. Volume 1: Aerodynamic methods and program users' guide

NASA Technical Reports Server (NTRS)

Axelson, J. A.

1977-01-01

The AEROX program estimates lift, induced-drag and pitching moments to high angles (typ. 60 deg) for wings and for wingbody combinations with or without an aft horizontal tail. Minimum drag coefficients are not estimated, but may be input for inclusion in the total aerodynamic parameters which are output in listed and plotted formats. The theory, users' guide, test cases, and program listing are presented.

Ramjets: Airframe Integration

DTIC Science & Technology

2010-09-01

nozzle • Brayton (or Joule) cycle: combustion at constant pressure at non-zero velocity The combustion process is modelled by means of adding heat to...against aerodynamic heating Aerodynamic heating calculations are based on: • Taylor -Maccoll method for compressible inviscid cone flow • Reynolds

The research progress on Hodograph Method of aerodynamic design at Tsinghua University

NASA Technical Reports Server (NTRS)

Chen, Zuoyi; Guo, Jingrong

1991-01-01

Progress in the use of the Hodograph method of aerodynamic design is discussed. It was found that there are some restricted conditions in the application of Hodograph design to transonic turbine and compressor cascades. The Hodograph method is suitable not only to the transonic turbine cascade but also to the transonic compressor cascade. The three dimensional Hodograph method will be developed after obtaining the basic equation for the three dimensional Hodograph method. As an example of the Hodograph method, the use of the method to design a transonic turbine and compressor cascade is discussed.

Aerodynamic Shape Optimization Using Hybridized Differential Evolution

NASA Technical Reports Server (NTRS)

Madavan, Nateri K.

2003-01-01

An aerodynamic shape optimization method that uses an evolutionary algorithm known at Differential Evolution (DE) in conjunction with various hybridization strategies is described. DE is a simple and robust evolutionary strategy that has been proven effective in determining the global optimum for several difficult optimization problems. Various hybridization strategies for DE are explored, including the use of neural networks as well as traditional local search methods. A Navier-Stokes solver is used to evaluate the various intermediate designs and provide inputs to the hybrid DE optimizer. The method is implemented on distributed parallel computers so that new designs can be obtained within reasonable turnaround times. Results are presented for the inverse design of a turbine airfoil from a modern jet engine. (The final paper will include at least one other aerodynamic design application). The capability of the method to search large design spaces and obtain the optimal airfoils in an automatic fashion is demonstrated.

Preliminary Structural Sensitivity Study of Hypersonic Inflatable Aerodynamic Decelerator Using Probabilistic Methods

NASA Technical Reports Server (NTRS)

Lyle, Karen H.

2014-01-01

Acceptance of new spacecraft structural architectures and concepts requires validated design methods to minimize the expense involved with technology validation via flighttesting. This paper explores the implementation of probabilistic methods in the sensitivity analysis of the structural response of a Hypersonic Inflatable Aerodynamic Decelerator (HIAD). HIAD architectures are attractive for spacecraft deceleration because they are lightweight, store compactly, and utilize the atmosphere to decelerate a spacecraft during re-entry. However, designers are hesitant to include these inflatable approaches for large payloads or spacecraft because of the lack of flight validation. In the example presented here, the structural parameters of an existing HIAD model have been varied to illustrate the design approach utilizing uncertainty-based methods. Surrogate models have been used to reduce computational expense several orders of magnitude. The suitability of the design is based on assessing variation in the resulting cone angle. The acceptable cone angle variation would rely on the aerodynamic requirements.

Development of computational methods for unsteady aerodynamics at the NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Yates, E. Carson, Jr.; Whitlow, Woodrow, Jr.

1987-01-01

The current scope, recent progress, and plans for research and development of computational methods for unsteady aerodynamics at the NASA Langley Research Center are reviewed. Both integral equations and finite difference methods for inviscid and viscous flows are discussed. Although the great bulk of the effort has focused on finite difference solution of the transonic small perturbation equation, the integral equation program is given primary emphasis here because it is less well known.

Development of computational methods for unsteady aerodynamics at the NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Yates, E. Carson, Jr.; Whitlow, Woodrow, Jr.

1987-01-01

The current scope, recent progress, and plans for research and development of computational methods for unsteady aerodynamics at the NASA Langley Research Center are reviewed. Both integral-equations and finite-difference method for inviscid and viscous flows are discussed. Although the great bulk of the effort has focused on finite-difference solution of the transonic small-perturbation equation, the integral-equation program is given primary emphasis here because it is less well known.

Calculation and Identification of the Aerodynamic Parameters for Small-Scaled Fixed-Wing UAVs.

PubMed

Shen, Jieliang; Su, Yan; Liang, Qing; Zhu, Xinhua

2018-01-13

The establishment of the Aircraft Dynamic Model(ADM) constitutes the prerequisite for the design of the navigation and control system, but the aerodynamic parameters in the model could not be readily obtained especially for small-scaled fixed-wing UAVs. In this paper, the procedure of computing the aerodynamic parameters is developed. All the longitudinal and lateral aerodynamic derivatives are firstly calculated through semi-empirical method based on the aerodynamics, rather than the wind tunnel tests or fluid dynamics software analysis. Secondly, the residuals of each derivative are proposed to be identified or estimated further via Extended Kalman Filter(EKF), with the observations of the attitude and velocity from the airborne integrated navigation system. Meanwhile, the observability of the targeted parameters is analyzed and strengthened through multiple maneuvers. Based on a small-scaled fixed-wing aircraft driven by propeller, the airborne sensors are chosen and the model of the actuators are constructed. Then, real flight tests are implemented to verify the calculation and identification process. Test results tell the rationality of the semi-empirical method and show the improvement of accuracy of ADM after the compensation of the parameters.

Calculation and Identification of the Aerodynamic Parameters for Small-Scaled Fixed-Wing UAVs

PubMed Central

Shen, Jieliang; Su, Yan; Liang, Qing; Zhu, Xinhua

2018-01-01

The establishment of the Aircraft Dynamic Model (ADM) constitutes the prerequisite for the design of the navigation and control system, but the aerodynamic parameters in the model could not be readily obtained especially for small-scaled fixed-wing UAVs. In this paper, the procedure of computing the aerodynamic parameters is developed. All the longitudinal and lateral aerodynamic derivatives are firstly calculated through semi-empirical method based on the aerodynamics, rather than the wind tunnel tests or fluid dynamics software analysis. Secondly, the residuals of each derivative are proposed to be identified or estimated further via Extended Kalman Filter (EKF), with the observations of the attitude and velocity from the airborne integrated navigation system. Meanwhile, the observability of the targeted parameters is analyzed and strengthened through multiple maneuvers. Based on a small-scaled fixed-wing aircraft driven by propeller, the airborne sensors are chosen and the model of the actuators are constructed. Then, real flight tests are implemented to verify the calculation and identification process. Test results tell the rationality of the semi-empirical method and show the improvement of accuracy of ADM after the compensation of the parameters. PMID:29342856

Aircraft Noise Prediction Program theoretical manual: Propeller aerodynamics and noise

NASA Technical Reports Server (NTRS)

Zorumski, W. E. (Editor); Weir, D. S. (Editor)

1986-01-01

The prediction sequence used in the aircraft noise prediction program (ANOPP) is described. The elements of the sequence are called program modules. The first group of modules analyzes the propeller geometry, the aerodynamics, including both potential and boundary-layer flow, the propeller performance, and the surface loading distribution. This group of modules is based entirely on aerodynamic strip theory. The next group of modules deals with the first group. Predictions of periodic thickness and loading noise are determined with time-domain methods. Broadband noise is predicted by a semiempirical method. Near-field predictions of fuselage surface pressrues include the effects of boundary layer refraction and scattering. Far-field predictions include atmospheric and ground effects.

Fluxes of Nitrous Acid (HONO) above an Agricultural Field Side near Paris

NASA Astrophysics Data System (ADS)

Laufs, S.; Cazaunau, M.; Stella, P.; Loubet, B.; Kurtenbach, R.; Cellier, P.; Mellouki, W.; Kleffmann, J.

2012-04-01

HONO is an important precursor of the OH radical, the detergent of the atmosphere. Field measurements show high diurnal HONO mixing ratios that cannot be explained by chemical models with known gas phase chemistry. Therefore, daytime sources of HONO are still under discussion. During the last decade many experimental investigation were performed to study heterogeneous production of HONO like the photo enhanced reduction of NO2 on humic acids or photolysis of HNO3 on surfaces. Recently, nitrite produced by bacteria, present in soil, was discussed as a source of HONO as well. In addition gas phase sources like the photolysis of nitrophenols, or the reaction of excited NO2 are discussed. Gradient measurements show high mixing ratios of HONO even above the boundary layer. However, beside intensive investigations on the sources of HONO, it is still an open question whether heterogeneous or gas phase sources are more important in the atmosphere. Flux measurements could represent a method to find the origin of missing sources of HONO. Until now instruments are not sensitive and fast enough to do Eddy correlation measurements for HONO. Alternatively, HONO fluxes are estimated by the Aerodynamic Gradient (AGM), or Relaxed Eddy Accumulation (REA) methods. Here we present HONO fluxes estimated by AGM and the LOPAP technique (Long Path Absorption Photometer) above an agricultural field in Grignon, Paris (48°51'N, 1°58'E). Fluxes during different seasons and different types of vegetations including bare soil will be presented and compared with chemical corrected fluxes of NO, NO2 and O3, or other parameters.

Study of Aerodynamic Drag Reduction on a Full-Scale Tractor-Trailer

DOT National Transportation Integrated Search

1976-04-01

Aerodynamic drag tests were performed on a tractor-trailer combination using the coast-down method on a smooth, nearly level runway. The tests included an investigation of drag reduction obtained with add-on devices that are commercially available or...

User's Manual for DuctE3D: A Program for 3D Euler Unsteady Aerodynamic and Aeroelastic Analysis of Ducted Fans

NASA Technical Reports Server (NTRS)

Srivastava, R.; Reddy, T. S. R.

1997-01-01

The program DuctE3D is used for steady or unsteady aerodynamic and aeroelastic analysis of ducted fans. This guide describes the input data required and the output files generated, in using DuctE3D. The analysis solves three dimensional unsteady, compressible Euler equations to obtain the aerodynamic forces. A normal mode structural analysis is used to obtain the aeroelastic equations, which are solved using either the time domain or the frequency domain solution method. Sample input and output files are included in this guide for steady aerodynamic analysis and aeroelastic analysis of an isolated fan row.

Measurements of Aerodynamic Characteristics of a 35 deg Sweptback NACA 65-009 Airfoil Model with 1/4-Chord Plain Flap by the NACA Wing-Flow Method

DTIC Science & Technology

1947-08-05

neral, the changein hin&- mount coefficientwith Gha,n@in angleof attackis practically zero for the first ‘j”of angle.of attack. At ,hl@ anglesof...aerodynamic chordat a Mach numberof O.$X)to 33 percentof themean aerodynamic chordat a Mach numberof J.lo, The smallrandomvariationsof . b NAOARM No...flap deflection,disre~ding tie apparentrandom scatter,movesrearwardgrad- from 65.percentman aergdynemic . chordat ,‘“M= 0.55 to & +~rcent mg

Special opportunities in helicopter aerodynamics

NASA Technical Reports Server (NTRS)

Mccroskey, W. J.

1983-01-01

Aerodynamic research relating to modern helicopters includes the study of three dimensional, unsteady, nonlinear flow fields. A selective review is made of some of the phenomenon that hamper the development of satisfactory engineering prediction techniques, but which provides a rich source of research opportunities: flow separations, compressibility effects, complex vortical wakes, and aerodynamic interference between components. Several examples of work in progress are given, including dynamic stall alleviation, the development of computational methods for transonic flow, rotor-wake predictions, and blade-vortex interactions.

The present status and the future of missile aerodynamics

NASA Technical Reports Server (NTRS)

Nielsen, Jack N.

1989-01-01

Recent developments in the state of the art in missile aerodynamics are reviewed. Among the subjects covered are: (1) Tri-service/NASA data base, (2) wing-body interference, (3) nonlinear controls, (4) hypersonic transition, (5) vortex interference, (6) airbreathers, supersonic inlets, (7) store separation problems, (8) correlation of missile data, (9) CFD codes for complete configurations, (10) engineering prediction methods, and (11) future configurations. Suggestions are made for future research and development to advance the state of the art of missile aerodynamics.

The present status and the future of missile aerodynamics

NASA Technical Reports Server (NTRS)

Nielsen, Jack N.

1988-01-01

Some recent developments in the state of the art in missile aerodynamics are reviewed. Among the subjects covered are: (1) tri-service/NASA data base, (2) wing-body interference, (3) nonlinear controls, (4) hypersonic transition, (5) vortex interference, (6) airbreathers, supersonic inlets, (7) store separation problems, (8) correlation of missile data, (9) CFD codes for complete configurations, (10) engineering prediction methods, and (11) future configurations. Suggestions are made for future research and development to advance the state of the art of missile aerodynamics.

Missile aerodynamics; Proceedings of the Conference, Monterey, CA, Oct. 31-Nov. 2, 1988

DOE Office of Scientific and Technical Information (OSTI.GOV)

Mendenhall, M.R.; Nixon, D.; Dillenius, M.F.E.

1989-01-01

The present conference discusses the development status of predictive capabilities for missile aerodynamic characteristics, the application of experimental techniques to missile-release problems, prospective high-performance missile designs, the use of lateral jet controls for missile guidance, and the integration of stores on modern tactical aircraft. Also discussed are semiempirical aerodynamic methods for preliminary design, high angle-of-attack behavior for an advanced missile, and the dynamic derivatives of missiles and fighter-type configurations at high angles-of-attack.

A modular approach to large-scale design optimization of aerospace systems

NASA Astrophysics Data System (ADS)

Hwang, John T.

Gradient-based optimization and the adjoint method form a synergistic combination that enables the efficient solution of large-scale optimization problems. Though the gradient-based approach struggles with non-smooth or multi-modal problems, the capability to efficiently optimize up to tens of thousands of design variables provides a valuable design tool for exploring complex tradeoffs and finding unintuitive designs. However, the widespread adoption of gradient-based optimization is limited by the implementation challenges for computing derivatives efficiently and accurately, particularly in multidisciplinary and shape design problems. This thesis addresses these difficulties in two ways. First, to deal with the heterogeneity and integration challenges of multidisciplinary problems, this thesis presents a computational modeling framework that solves multidisciplinary systems and computes their derivatives in a semi-automated fashion. This framework is built upon a new mathematical formulation developed in this thesis that expresses any computational model as a system of algebraic equations and unifies all methods for computing derivatives using a single equation. The framework is applied to two engineering problems: the optimization of a nanosatellite with 7 disciplines and over 25,000 design variables; and simultaneous allocation and mission optimization for commercial aircraft involving 330 design variables, 12 of which are integer variables handled using the branch-and-bound method. In both cases, the framework makes large-scale optimization possible by reducing the implementation effort and code complexity. The second half of this thesis presents a differentiable parametrization of aircraft geometries and structures for high-fidelity shape optimization. Existing geometry parametrizations are not differentiable, or they are limited in the types of shape changes they allow. This is addressed by a novel parametrization that smoothly interpolates aircraft components, providing differentiability. An unstructured quadrilateral mesh generation algorithm is also developed to automate the creation of detailed meshes for aircraft structures, and a mesh convergence study is performed to verify that the quality of the mesh is maintained as it is refined. As a demonstration, high-fidelity aerostructural analysis is performed for two unconventional configurations with detailed structures included, and aerodynamic shape optimization is applied to the truss-braced wing, which finds and eliminates a shock in the region bounded by the struts and the wing.

Steady, Oscillatory, and Unsteady Subsonic and Supersonic Aerodynamics, production version (SOUSSA-P 1.1). Volume 1: Theoretical manual. [Green function

NASA Technical Reports Server (NTRS)

Morino, L.

1980-01-01

Recent developments of the Green's function method and the computer program SOUSSA (Steady, Oscillatory, and Unsteady Subsonic and Supersonic Aerodynamics) are reviewed and summarized. Applying the Green's function method to the fully unsteady (transient) potential equation yields an integro-differential-delay equation. With spatial discretization by the finite-element method, this equation is approximated by a set of differential-delay equations in time. Time solution by Laplace transform yields a matrix relating the velocity potential to the normal wash. Premultiplying and postmultiplying by the matrices relating generalized forces to the potential and the normal wash to the generalized coordinates one obtains the matrix of the generalized aerodynamic forces. The frequency and mode-shape dependence of this matrix makes the program SOUSSA useful for multiple frequency and repeated mode-shape evaluations.

Nonlinear programming extensions to rational function approximation methods for unsteady aerodynamic forces

NASA Technical Reports Server (NTRS)

Tiffany, Sherwood H.; Adams, William M., Jr.

1988-01-01

The approximation of unsteady generalized aerodynamic forces in the equations of motion of a flexible aircraft are discussed. Two methods of formulating these approximations are extended to include the same flexibility in constraining the approximations and the same methodology in optimizing nonlinear parameters as another currently used extended least-squares method. Optimal selection of nonlinear parameters is made in each of the three methods by use of the same nonlinear, nongradient optimizer. The objective of the nonlinear optimization is to obtain rational approximations to the unsteady aerodynamics whose state-space realization is lower order than that required when no optimization of the nonlinear terms is performed. The free linear parameters are determined using the least-squares matrix techniques of a Lagrange multiplier formulation of an objective function which incorporates selected linear equality constraints. State-space mathematical models resulting from different approaches are described and results are presented that show comparative evaluations from application of each of the extended methods to a numerical example.

Integration of a supersonic unsteady aerodynamic code into the NASA FASTEX system

NASA Technical Reports Server (NTRS)

Appa, Kari; Smith, Michael J. C.

1987-01-01

A supersonic unsteady aerodynamic loads prediction method based on the constant pressure method was integrated into the NASA FASTEX system. The updated FASTEX code can be employed for aeroelastic analyses in subsonic and supersonic flow regimes. A brief description of the supersonic constant pressure panel method, as applied to lifting surfaces and body configurations, is followed by a documentation of updates required to incorporate this method in the FASTEX code. Test cases showing correlations of predicted pressure distributions, flutter solutions, and stability derivatives with available data are reported.

Soaring with Science.

ERIC Educational Resources Information Center

Tamarkin, Cali; Bourne, Barbara

1995-01-01

Presents an integrated unit on aerodynamics for a diverse group of 4th grade students. Students work with readily available, inexpensive materials to devise methods to answer questions about aerodynamics. High-interest activities include balloon rocketry and paper-airplane design and address visual, auditory, and kinesthetic learning styles. The…

Advanced prediction technique for the low speed aerodynamics of V/STOL aircraft. Volume 1: Technical discussion

NASA Technical Reports Server (NTRS)

Beatty, T. D.; Worthey, M. K.

1984-01-01

The V/STOL Aircraft Propulsive Effects (VAPE) computerized prediction method is evaluated. The program analyzes viscous effects, various jet, inlet, and Short TakeOff and Landing (STOL) models, and examines the aerodynamic configurations of V/STOL aircraft.

Computer tomography of flows external to test models

NASA Technical Reports Server (NTRS)

Prikryl, I.; Vest, C. M.

1982-01-01

Computer tomographic techniques for reconstruction of three-dimensional aerodynamic density fields, from interferograms recorded from several different viewing directions were studied. Emphasis is on the case in which an opaque object such as a test model in a wind tunnel obscures significant regions of the interferograms (projection data). A method called the Iterative Convolution Method (ICM), existing methods in which the field is represented by a series expansions, and analysis of real experimental data in the form of aerodynamic interferograms are discussed.

Analysis of Aerodynamic Load of LSU-03 (LAPAN Surveillance UAV-03) Propeller

NASA Astrophysics Data System (ADS)

Rahmadi Nuranto, Awang; Jamaludin Fitroh, Ahmad; Syamsudin, Hendri

2018-04-01

The existing propeller of the LSU-03 aircraft is made of wood. To improve structural strength and obtain better mechanical properties, the propeller will be redesigned usingcomposite materials. It is necessary to simulate and analyze the design load. This research paper explainsthe simulation and analysis of aerodynamic load prior to structural design phase of composite propeller. Aerodynamic load calculations are performed using both the Blade Element Theory(BET) and the Computational Fluid Dynamic (CFD)simulation. The result of both methods show a close agreement, the different thrust forces is only 1.2 and 4.1% for two type mesh. Thus the distribution of aerodynamic loads along the surface of the propeller blades of the 3-D CFD simulation results are considered valid and ready to design the composite structure. TheCFD results is directly imported to the structure model using the Direct Import CFD / One-Way Fluid Structure Interaction (FSI) method. Design load of propeller is chosen at the flight condition at speed of 20 km/h at 7000 rpm.

Roll Damping Derivatives from Generalized Lifting-Surface Theory and Wind Tunnel Forced-Oscillation Tests

NASA Technical Reports Server (NTRS)

Pototzky, Anthony S; Murphy, Patrick C.

2014-01-01

Improving aerodynamic models for adverse loss-of-control conditions in flight is an area being researched under the NASA Aviation Safety Program. Aerodynamic models appropriate for loss of control conditions require a more general mathematical representation to predict nonlinear unsteady behaviors. As more general aerodynamic models are studied that include nonlinear higher order effects, the possibility of measurements that confound aerodynamic and structural responses are probable. In this study an initial step is taken to look at including structural flexibility in analysis of rigid-body forced-oscillation testing that accounts for dynamic rig, sting and balance flexibility. Because of the significant testing required and associated costs in a general study, it makes sense to capitalize on low cost analytical methods where possible, especially where structural flexibility can be accounted for by a low cost method. This paper provides an initial look at using linear lifting surface theory applied to rigid-body aircraft roll forced-oscillation tests.

Feedback tracking control for dynamic morphing of piezocomposite actuated flexible wings

NASA Astrophysics Data System (ADS)

Wang, Xiaoming; Zhou, Wenya; Wu, Zhigang

2018-03-01

Aerodynamic properties of flexible wings can be improved via shape morphing using piezocomposite materials. Dynamic shape control of flexible wings is investigated in this study by considering the interactions between structural dynamics, unsteady aerodynamics and piezo-actuations. A novel antisymmetric angle-ply bimorph configuration of piezocomposite actuators is presented to realize coupled bending-torsional shape control. The active aeroelastic model is derived using finite element method and Theodorsen unsteady aerodynamic loads. A time-varying linear quadratic Gaussian (LQG) tracking control system is designed to enhance aerodynamic lift with pre-defined trajectories. Proof-of-concept simulations of static and dynamic shape control are presented for a scaled high-aspect-ratio wing model. Vibrations of the wing and fluctuations in aerodynamic forces are caused by using the static voltages directly in dynamic shape control. The lift response has tracked the trajectories well with favorable dynamic morphing performance via feedback tracking control.

High-angle-of-attack aerodynamics - Lessons learned

NASA Technical Reports Server (NTRS)

Chambers, J. R.

1986-01-01

Recently, the military and civil technical communities have undertaken numerous studies of the high angle-of-attack aerodynamic characteristics of advanced airplane and missile configurations. The method of approach and the design methodology employed have necessarily been experimental and exploratory in nature, due to the complex nature of separated flows. However, despite the relatively poor definition of many of the key aerodynamic phenomena involved for high-alpha conditions, some generic guidelines for design consideration have been identified. The present paper summarizes some of the more important lessons learned in the area of high angle-of-attack aerodynamics with examples of a number of key concepts and with particular emphasis on high-alpha stability and control characteristics of high performance aircraft. Topics covered in the discussion include the impact of design evolution, forebody flows, control of separated flows, configuration effects, aerodynamic controls, wind-tunnel flight correlation, and recent NASA research activities.

Single- and Multiple-Objective Optimization with Differential Evolution and Neural Networks

NASA Technical Reports Server (NTRS)

Rai, Man Mohan

2006-01-01

Genetic and evolutionary algorithms have been applied to solve numerous problems in engineering design where they have been used primarily as optimization procedures. These methods have an advantage over conventional gradient-based search procedures became they are capable of finding global optima of multi-modal functions and searching design spaces with disjoint feasible regions. They are also robust in the presence of noisy data. Another desirable feature of these methods is that they can efficiently use distributed and parallel computing resources since multiple function evaluations (flow simulations in aerodynamics design) can be performed simultaneously and independently on ultiple processors. For these reasons genetic and evolutionary algorithms are being used more frequently in design optimization. Examples include airfoil and wing design and compressor and turbine airfoil design. They are also finding increasing use in multiple-objective and multidisciplinary optimization. This lecture will focus on an evolutionary method that is a relatively new member to the general class of evolutionary methods called differential evolution (DE). This method is easy to use and program and it requires relatively few user-specified constants. These constants are easily determined for a wide class of problems. Fine-tuning the constants will off course yield the solution to the optimization problem at hand more rapidly. DE can be efficiently implemented on parallel computers and can be used for continuous, discrete and mixed discrete/continuous optimization problems. It does not require the objective function to be continuous and is noise tolerant. DE and applications to single and multiple-objective optimization will be included in the presentation and lecture notes. A method for aerodynamic design optimization that is based on neural networks will also be included as a part of this lecture. The method offers advantages over traditional optimization methods. It is more flexible than other methods in dealing with design in the context of both steady and unsteady flows, partial and complete data sets, combined experimental and numerical data, inclusion of various constraints and rules of thumb, and other issues that characterize the aerodynamic design process. Neural networks provide a natural framework within which a succession of numerical solutions of increasing fidelity, incorporating more realistic flow physics, can be represented and utilized for optimization. Neural networks also offer an excellent framework for multiple-objective and multi-disciplinary design optimization. Simulation tools from various disciplines can be integrated within this framework and rapid trade-off studies involving one or many disciplines can be performed. The prospect of combining neural network based optimization methods and evolutionary algorithms to obtain a hybrid method with the best properties of both methods will be included in this presentation. Achieving solution diversity and accurate convergence to the exact Pareto front in multiple objective optimization usually requires a significant computational effort with evolutionary algorithms. In this lecture we will also explore the possibility of using neural networks to obtain estimates of the Pareto optimal front using non-dominated solutions generated by DE as training data. Neural network estimators have the potential advantage of reducing the number of function evaluations required to obtain solution accuracy and diversity, thus reducing cost to design.

A practical model for pressure probe system response estimation (with review of existing models)

NASA Astrophysics Data System (ADS)

Hall, B. F.; Povey, T.

2018-04-01

The accurate estimation of the unsteady response (bandwidth) of pneumatic pressure probe systems (probe, line and transducer volume) is a common practical problem encountered in the design of aerodynamic experiments. Understanding the bandwidth of the probe system is necessary to capture unsteady flow features accurately. Where traversing probes are used, the desired traverse speed and spatial gradients in the flow dictate the minimum probe system bandwidth required to resolve the flow. Existing approaches for bandwidth estimation are either complex or inaccurate in implementation, so probes are often designed based on experience. Where probe system bandwidth is characterized, it is often done experimentally, requiring careful experimental set-up and analysis. There is a need for a relatively simple but accurate model for estimation of probe system bandwidth. A new model is presented for the accurate estimation of pressure probe bandwidth for simple probes commonly used in wind tunnel environments; experimental validation is provided. An additional, simple graphical method for air is included for convenience.

Application of a laser interferometer skin-friction meter in complex flows

NASA Technical Reports Server (NTRS)

Monson, D. J.; Driver, D. M.; Szodruch, J.

1981-01-01

The application of a nonintrusive laser-interferometer skin-friction meter, which measures skin friction with a remotely located laser interferometer that monitors the thickness change of a thin oil film, is extended both experimentally and theoretically to several complex wind-tunnel flows. These include two-dimensional seperated and reattached subsonic flows with large pressure and shear gradients, and two and three-dimensional supersonic flows at high Reynolds number, which include variable wall temperatures and cross-flows. In addition, it is found that the instrument can provide an accurate location of the mean reattachment length for separated flows. Results show that levels up to 120 N/sq m, or 40 times higher than previous tests, can be obtained, despite encountering some limits to the method for very high skin-friction levels. It is concluded that these results establish the utility of this instrument for measuring skin friction in a wide variety of flows of interest in aerodynamic testing.

Aerodynamic calculational methods for curved-blade Darrieus VAWT WECS

NASA Astrophysics Data System (ADS)

Templin, R. J.

1985-03-01

Calculation of aerodynamic performance and load distributions for curved-blade wind turbines is discussed. Double multiple stream tube theory, and the uncertainties that remain in further developing adequate methods are considered. The lack of relevant airfoil data at high Reynolds numbers and high angles of attack, and doubts concerning the accuracy of models of dynamic stall are underlined. Wind tunnel tests of blade airbrake configurations are summarized.

Aerodynamic levitation, supercooled liquids and glass formation

DOE PAGES

Benmore, C. J.; Weber, J. K. R.

2017-05-04

Containerless processing or ‘levitation’ is a valuable tool for the synthesis and characterization of materials, particularly at extreme temperatures and under non-equilibrium conditions. The method enables formation of novel glasses, amorphous phases, and metastable crystalline forms that are not easily accessed when nucleation and growth can readily occur at a container interface. Removing the container enables the use of a wide variety of process atmospheres to modify a materials structure and properties. In the past decade levitation methods, including acoustic, aerodynamic, electromagnetic, and electrostatic, have become well established sample environments at X-ray synchrotron and neutron sources. This article briefly reviewsmore » the methods and then focuses on the application of aerodynamic levitation to synthesize and study new materials. This is presented in conjunction with non-contact probes used to investigate the atomic structure and to measure the properties of materials at extreme temperatures. The use of aerodynamic levitation in research using small and wide-angle X-ray diffraction, XANES, and neutron scattering are discussed in the context of technique development. The use of the containerless methods to investigate thermophysical properties is also considered. We argue that structural motifs and in the liquid state can potentially lead to the fabrication of materials, whose properties would differ substantially from their well known crystalline forms.« less

Aerodynamic Design Study of Advanced Multistage Axial Compressor

NASA Technical Reports Server (NTRS)

Larosiliere, Louis M.; Wood, Jerry R.; Hathaway, Michael D.; Medd, Adam J.; Dang, Thong Q.

2002-01-01

As a direct response to the need for further performance gains from current multistage axial compressors, an investigation of advanced aerodynamic design concepts that will lead to compact, high-efficiency, and wide-operability configurations is being pursued. Part I of this report describes the projected level of technical advancement relative to the state of the art and quantifies it in terms of basic aerodynamic technology elements of current design systems. A rational enhancement of these elements is shown to lead to a substantial expansion of the design and operability space. Aerodynamic design considerations for a four-stage core compressor intended to serve as a vehicle to develop, integrate, and demonstrate aerotechnology advancements are discussed. This design is biased toward high efficiency at high loading. Three-dimensional blading and spanwise tailoring of vector diagrams guided by computational fluid dynamics (CFD) are used to manage the aerodynamics of the high-loaded endwall regions. Certain deleterious flow features, such as leakage-vortex-dominated endwall flow and strong shock-boundary-layer interactions, were identified and targeted for improvement. However, the preliminary results were encouraging and the front two stages were extracted for further aerodynamic trimming using a three-dimensional inverse design method described in part II of this report. The benefits of the inverse design method are illustrated by developing an appropriate pressure-loading strategy for transonic blading and applying it to reblade the rotors in the front two stages of the four-stage configuration. Multistage CFD simulations based on the average passage formulation indicated an overall efficiency potential far exceeding current practice for the front two stages. Results of the CFD simulation at the aerodynamic design point are interrogated to identify areas requiring additional development. In spite of the significantly higher aerodynamic loadings, advanced CFD-based tools were able to effectively guide the design of a very efficient axial compressor under state-of-the-art aeromechanical constraints.

Evaluation of a CFD Method for Aerodynamic Database Development using the Hyper-X Stack Configuration

NASA Technical Reports Server (NTRS)

Parikh, Paresh; Engelund, Walter; Armand, Sasan; Bittner, Robert

2004-01-01

A computational fluid dynamic (CFD) study is performed on the Hyper-X (X-43A) Launch Vehicle stack configuration in support of the aerodynamic database generation in the transonic to hypersonic flow regime. The main aim of the study is the evaluation of a CFD method that can be used to support aerodynamic database development for similar future configurations. The CFD method uses the NASA Langley Research Center developed TetrUSS software, which is based on tetrahedral, unstructured grids. The Navier-Stokes computational method is first evaluated against a set of wind tunnel test data to gain confidence in the code s application to hypersonic Mach number flows. The evaluation includes comparison of the longitudinal stability derivatives on the complete stack configuration (which includes the X-43A/Hyper-X Research Vehicle, the launch vehicle and an adapter connecting the two), detailed surface pressure distributions at selected locations on the stack body and component (rudder, elevons) forces and moments. The CFD method is further used to predict the stack aerodynamic performance at flow conditions where no experimental data is available as well as for component loads for mechanical design and aero-elastic analyses. An excellent match between the computed and the test data over a range of flow conditions provides a computational tool that may be used for future similar hypersonic configurations with confidence.

Identification of Linear and Nonlinear Aerodynamic Impulse Responses Using Digital Filter Techniques

NASA Technical Reports Server (NTRS)

Silva, Walter A.

1997-01-01

This paper discusses the mathematical existence and the numerically-correct identification of linear and nonlinear aerodynamic impulse response functions. Differences between continuous-time and discrete-time system theories, which permit the identification and efficient use of these functions, will be detailed. Important input/output definitions and the concept of linear and nonlinear systems with memory will also be discussed. It will be shown that indicial (step or steady) responses (such as Wagner's function), forced harmonic responses (such as Theodorsen's function or those from doublet lattice theory), and responses to random inputs (such as gusts) can all be obtained from an aerodynamic impulse response function. This paper establishes the aerodynamic impulse response function as the most fundamental, and, therefore, the most computationally efficient, aerodynamic function that can be extracted from any given discrete-time, aerodynamic system. The results presented in this paper help to unify the understanding of classical two-dimensional continuous-time theories with modern three-dimensional, discrete-time theories. First, the method is applied to the nonlinear viscous Burger's equation as an example. Next the method is applied to a three-dimensional aeroelastic model using the CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) code and then to a two-dimensional model using the CFL3D Navier-Stokes code. Comparisons of accuracy and computational cost savings are presented. Because of its mathematical generality, an important attribute of this methodology is that it is applicable to a wide range of nonlinear, discrete-time problems.

Identification of Linear and Nonlinear Aerodynamic Impulse Responses Using Digital Filter Techniques

NASA Technical Reports Server (NTRS)

Silva, Walter A.

1997-01-01

This paper discusses the mathematical existence and the numerically-correct identification of linear and nonlinear aerodynamic impulse response functions. Differences between continuous-time and discrete-time system theories, which permit the identification and efficient use of these functions, will be detailed. Important input/output definitions and the concept of linear and nonlinear systems with memory will also be discussed. It will be shown that indicial (step or steady) responses (such as Wagner's function), forced harmonic responses (such as Tbeodorsen's function or those from doublet lattice theory), and responses to random inputs (such as gusts) can all be obtained from an aerodynamic impulse response function. This paper establishes the aerodynamic impulse response function as the most fundamental, and, therefore, the most computationally efficient, aerodynamic function that can be extracted from any given discrete-time, aerodynamic system. The results presented in this paper help to unify the understanding of classical two-dimensional continuous-time theories with modem three-dimensional, discrete-time theories. First, the method is applied to the nonlinear viscous Burger's equation as an example. Next the method is applied to a three-dimensional aeroelastic model using the CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) code and then to a two-dimensional model using the CFL3D Navier-Stokes code. Comparisons of accuracy and computational cost savings are presented. Because of its mathematical generality, an important attribute of this methodology is that it is applicable to a wide range of nonlinear, discrete-time problems.

Numerical study on the aerodynamic characteristics of both static and flapping wing with attachments

NASA Astrophysics Data System (ADS)

Xie, Lingwang; Zhang, Xingwei; Luo, Pan; Huang, Panpan

2017-10-01

The purpose of this paper is to investigate the aerodynamic mechanism of airfoils under different icing situations which are different icing type, different icing time, and different icing position. Numerical simulation is carried out by using the finite volume method for both static and flapping airfoils, when Reynolds number is kept at 135000. The difference of aerodynamic performance between the airfoil with attachments and without attachments are be investigated by comparing the force coefficients, lift-to-drag ratios and flow field contour. The present simulations reveal that some influences of attachment are similar in the static airfoil and the flapping airfoil. Specifically, the airfoil with the attachment derived from glaze ice type causes the worse aerodynamic performance than that derived from rime ice type. The longer the icing time, the greater influence of aerodynamic performance the attachment causes. The attachments on the leading-edge have the greater influence of aerodynamic performance than other positions. Moreover, there are little differences between the static airfoil and the flapping airfoil. Compared with the static airfoil, the flapping airfoil which attachment located on the trailing edge causes a worse aerodynamic performance. Both attachments derived from rime ice type and glaze ice type all will deteriorate the aerodynamic performance of the asymmetrical airfoils. Present work provides the systematic and comprehensive study about icing blade which is conducive to the development of the wind power generation technology.

Aerodynamic database development of the ESA intermediate experimental vehicle

NASA Astrophysics Data System (ADS)

Pezzella, Giuseppe; Marino, Giuliano; Rufolo, Giuseppe C.

2014-01-01

This work deals with the aerodynamic database development of the Intermediate Experiment Vehicle. The aerodynamic analysis, carried out for the whole flight scenario, relies on computational fluid dynamics, wind tunnel test, and engineering-based design data generated during the project phases, from rarefied flow conditions, to hypersonic continuum flow up to reach subsonic speeds regime. Therefore, the vehicle aerodynamic database covers the range of Mach number, angle of attack, sideslip and control surface deflections foreseen for the vehicle nominal re-entry. In particular, the databasing activities are developed in the light of build-up approach. This means that all aerodynamic force and moment coefficients are provided by means of a linear summation over certain number of incremental contributions such as, for example, effect of sideslip angle, aerodynamic control surface effectiveness, etc. Each force and moment coefficient is treated separately and appropriate equation is provided, in which all the pertinent contributions for obtaining the total coefficient for any selected flight conditions appear. To this aim, all the available numerical and experimental aerodynamic data are gathered in order to explicit the functional dependencies from each aerodynamic model addend through polynomial expressions obtained with the least squares method. These polynomials are function of the primary variable that drives the phenomenon whereas secondary dependencies are introduced directly into its unknown coefficients which are determined by means of best-fitting algorithms.

A Quasi-Steady Flexible Launch Vehicle Stability Analysis Using Steady CFD with Unsteady Aerodynamic Enhancement

NASA Technical Reports Server (NTRS)

Bartels, Robert E.

2011-01-01

Launch vehicles frequently experience a reduced stability margin through the transonic Mach number range. This reduced stability margin is caused by an undamping of the aerodynamics in one of the lower frequency flexible or rigid body modes. Analysis of the behavior of a flexible vehicle is routinely performed with quasi-steady aerodynamic lineloads derived from steady rigid computational fluid dynamics (CFD). However, a quasi-steady aeroelastic stability analysis can be unconservative at the critical Mach numbers where experiment or unsteady computational aeroelastic (CAE) analysis show a reduced or even negative aerodynamic damping. This paper will present a method of enhancing the quasi-steady aeroelastic stability analysis of a launch vehicle with unsteady aerodynamics. The enhanced formulation uses unsteady CFD to compute the response of selected lower frequency modes. The response is contained in a time history of the vehicle lineloads. A proper orthogonal decomposition of the unsteady aerodynamic lineload response is used to reduce the scale of data volume and system identification is used to derive the aerodynamic stiffness, damping and mass matrices. The results of the enhanced quasi-static aeroelastic stability analysis are compared with the damping and frequency computed from unsteady CAE analysis and from a quasi-steady analysis. The results show that incorporating unsteady aerodynamics in this way brings the enhanced quasi-steady aeroelastic stability analysis into close agreement with the unsteady CAE analysis.

Using High Resolution Design Spaces for Aerodynamic Shape Optimization Under Uncertainty

NASA Technical Reports Server (NTRS)

Li, Wu; Padula, Sharon

2004-01-01

This paper explains why high resolution design spaces encourage traditional airfoil optimization algorithms to generate noisy shape modifications, which lead to inaccurate linear predictions of aerodynamic coefficients and potential failure of descent methods. By using auxiliary drag constraints for a simultaneous drag reduction at all design points and the least shape distortion to achieve the targeted drag reduction, an improved algorithm generates relatively smooth optimal airfoils with no severe off-design performance degradation over a range of flight conditions, in high resolution design spaces parameterized by cubic B-spline functions. Simulation results using FUN2D in Euler flows are included to show the capability of the robust aerodynamic shape optimization method over a range of flight conditions.

Estimation et validation des derivees de stabilite et controle du modele dynamique non-lineaire d'un drone a voilure fixe

NASA Astrophysics Data System (ADS)

Courchesne, Samuel

Knowledge of the dynamic characteristics of a fixed-wing UAV is necessary to design flight control laws and to conceive a high quality flight simulator. The basic features of a flight mechanic model include the properties of mass, inertia and major aerodynamic terms. They respond to a complex process involving various numerical analysis techniques and experimental procedures. This thesis focuses on the analysis of estimation techniques applied to estimate problems of stability and control derivatives from flight test data provided by an experimental UAV. To achieve this objective, a modern identification methodology (Quad-M) is used to coordinate the processing tasks from multidisciplinary fields, such as parameter estimation modeling, instrumentation, the definition of flight maneuvers and validation. The system under study is a non-linear model with six degrees of freedom with a linear aerodynamic model. The time domain techniques are used for identification of the drone. The first technique, the equation error method is used to determine the structure of the aerodynamic model. Thereafter, the output error method and filter error method are used to estimate the aerodynamic coefficients values. The Matlab scripts for estimating the parameters obtained from the American Institute of Aeronautics and Astronautics (AIAA) are used and modified as necessary to achieve the desired results. A commendable effort in this part of research is devoted to the design of experiments. This includes an awareness of the system data acquisition onboard and the definition of flight maneuvers. The flight tests were conducted under stable flight conditions and with low atmospheric disturbance. Nevertheless, the identification results showed that the filter error method is most effective for estimating the parameters of the drone due to the presence of process noise and measurement. The aerodynamic coefficients are validated using a numerical analysis of the vortex method. In addition, a simulation model incorporating the estimated parameters is used to compare the behavior of states measured. Finally, a good correspondence between the results is demonstrated despite a limited number of flight data. Keywords: drone, identification, estimation, nonlinear, flight test, system, aerodynamic coefficient.

Configuration Aerodynamics: Past - Present - Future

NASA Technical Reports Server (NTRS)

Wood, Richard M.; Agrawal, Shreekant; Bencze, Daniel P.; Kulfan, Robert M.; Wilson, Douglas L.

1999-01-01

The Configuration Aerodynamics (CA) element of the High Speed Research (HSR) program is managed by a joint NASA and Industry team, referred to as the Technology Integration Development (ITD) team. This team is responsible for the development of a broad range of technologies for improved aerodynamic performance and stability and control characteristics at subsonic to supersonic flight conditions. These objectives are pursued through the aggressive use of advanced experimental test techniques and state of the art computational methods. As the HSR program matures and transitions into the next phase the objectives of the Configuration Aerodynamics ITD are being refined to address the drag reduction needs and stability and control requirements of High Speed Civil Transport (HSCT) aircraft. In addition, the experimental and computational tools are being refined and improved to meet these challenges. The presentation will review the work performed within the Configuration Aerodynamics element in 1994 and 1995 and then discuss the plans for the 1996-1998 time period. The final portion of the presentation will review several observations of the HSR program and the design activity within Configuration Aerodynamics.

Influence of Different Diffuser Angle on Sedan's Aerodynamic Characteristics

NASA Astrophysics Data System (ADS)

Hu, Xingjun; Zhang, Rui; Ye, Jian; Yan, Xu; Zhao, Zhiming

The aerodynamic characteristics have a great influence on the fuel economics and the steering stability of a high speed vehicle. The underbody rear diffuser is one of important aerodynamic add-on devices. The parameters of the diffuser, including the diffuser angle, the number and the shape of separators, the shape of the end plate and etc, will affect the underbody flow and the wake. Here, just the influence of the diffuser angle was investigated without separator and the end plate. The method of Computational Fluid Dynamics was adopted to study the aerodynamic characteristics of a simplified sedan with a different diffuser angle respectively. The diffuser angle was set to 0°, 3°, 6°, 9.8° and 12° respectively. The diffuser angle of the original model is 9.8°. The conclusions were drawn that when the diffuser angle increases, the underbody flow and especially the wake change greatly and the pressure change correspondingly; as a result, the total aerodynamic drag coefficients of car first decrease and then increases, while the total aerodynamic lift coefficients decrease.

Aerodynamic Shape Optimization of a Dual-Stream Supersonic Plug Nozzle

NASA Technical Reports Server (NTRS)

Heath, Christopher M.; Gray, Justin S.; Park, Michael A.; Nielsen, Eric J.; Carlson, Jan-Renee

2015-01-01

Aerodynamic shape optimization was performed on an isolated axisymmetric plug nozzle sized for a supersonic business jet. The dual-stream concept was tailored to attenuate nearfield pressure disturbances without compromising nozzle performance. Adjoint-based anisotropic mesh refinement was applied to resolve nearfield compression and expansion features in the baseline viscous grid. Deformed versions of the adapted grid were used for subsequent adjoint-driven shape optimization. For design, a nonlinear gradient-based optimizer was coupled to the discrete adjoint formulation of the Reynolds-averaged Navier- Stokes equations. All nozzle surfaces were parameterized using 3rd order B-spline interpolants and perturbed axisymmetrically via free-form deformation. Geometry deformations were performed using 20 design variables shared between the outer cowl, shroud and centerbody nozzle surfaces. Interior volume grid deformation during design was accomplished using linear elastic mesh morphing. The nozzle optimization was performed at a design cruise speed of Mach 1.6, assuming core and bypass pressure ratios of 6.19 and 3.24, respectively. Ambient flight conditions at design were commensurate with 45,000-ft standard day atmosphere.

Flight measurements of surface pressures on a flexible supercritical research wing

NASA Technical Reports Server (NTRS)

Eckstrom, C. V.

1985-01-01

A flexible supercritical research wing, designated as ARW-1, was flight-tested as part of the NASA Drones for Aerodynamic and Structural Testing Program. Aerodynamic loads, in the form of wing surface pressure measurements, were obtained during flights at altitudes of 15,000, 20,000, and 25,000 feet at Mach numbers from 0.70 to 0.91. Surface pressure coefficients determined from pressure measurements at 80 orifice locations are presented individually as nearly continuous functions of angle of attack for constant values of Mach number. The surface pressure coefficients are also presented individually as a function of Mach number for an angle of attack of 2.0 deg. The nearly continuous values of the pressure coefficient clearly show details of the pressure gradient, which occurred in a rather narrow Mach number range. The effects of changes in angle of attack, Mach number, and dynamic pressure are also shown by chordwise pressure distributions for the range of test conditions experienced. Reynolds numbers for the tests ranged from 5.7 to 8.4 x 1,000,000.

Wind-Tunnel Balance Characterization for Hypersonic Research Applications

NASA Technical Reports Server (NTRS)

Lynn, Keith C.; Commo, Sean A.; Parker, Peter A.

2012-01-01

Wind-tunnel research was recently conducted at the NASA Langley Research Center s 31-Inch Mach 10 Hypersonic Facility in support of the Mars Science Laboratory s aerodynamic program. Researchers were interested in understanding the interaction between the freestream flow and the reaction control system onboard the entry vehicle. A five-component balance, designed for hypersonic testing with pressurized flow-through capability, was used. In addition to the aerodynamic forces, the balance was exposed to both thermal gradients and varying internal cavity pressures. Historically, the effect of these environmental conditions on the response of the balance have not been fully characterized due to the limitations in the calibration facilities. Through statistical design of experiments, thermal and pressure effects were strategically and efficiently integrated into the calibration of the balance. As a result of this new approach, researchers were able to use the balance continuously throughout the wide range of temperatures and pressures and obtain real-time results. Although this work focused on a specific application, the methodology shown can be applied more generally to any force measurement system calibration.

Preferred negative geotactic orientation in mobile cells: Tetrahymena results.

PubMed Central

Noever, D A; Cronise, R; Matsos, H C

1994-01-01

For the protozoan species Tetrahymena a series of airplane experiments are reported, which varied gravity as an active laboratory parameter and tested for corresponding changes in geotaxic orientation of single cells. The airplane achieved alternating periods of low (0.01 g) and high (1.8 g; g = 980 cm/s) gravity by flying repeated Keplerian parabolas. The experimental design was undertaken to clearly distinguish gravity from competing aerodynamic and chemical gradients. In this way, each culture served as its own control, with gravity level alone determining the orientational changes. On average, 6.3% of the Tetrahymena oriented vertically in low gravity, while 27% oriented vertically in high-gravity phases. Simplified physical models are explored for describing these cell trajectories as a function of gravity, aerodynamic drag, and lift. The notable effect of gravity on turning behavior is emphasized as the biophysical cause of the observed negative geotaxis in Tetrahymena. A fundamental investigation of the biological gravity receptor (if it exists) and improved modeling for vertical migration in important types of ocean plankton motivate the present research. Images FIGURE 1 PMID:7858146

Impact of implementation of spaceborne lidar-retrieved canopy height in the WRF model

NASA Astrophysics Data System (ADS)

Lee, Junhong; Hong, Jinkyu

2017-04-01

Canopy height is closely related to biomass and aerodynamic properties, which regulate turbulent transfer of energy and mass at the soil-vegetation-atmosphere continuum. However, this key information has been prescribed as a constant value in a fixed plant functional type in atmospheric models. This presentation reports impacts of using realistic forest canopy height, retrieved from spaceborne LiDAR, on regional climate simulation in the Weather Research and Forecasting (WRF) model's land surface model. Numerical simulations were conducted over the Amazon Basin and East Asia during summer season. Over these regions, the LiDAR-retrieved canopy heights were higher than the default values used in the WRF,which are dependent only on plant functional type. By modifying roughness length and zero-plane displacement height, the change of canopy height resulted in changes in surface energy balance by regulating aerodynamic conductances and vertical temperature gradient, thus modifying the lifting condensation level and equivalent potential temperature in the atmospheric boundary layer. Our analysis also showed that the WRF model better reproduced the observed precipitation when LiDAR-retrieved canopy height was used over the Amazon Basin.

Implementation of spaceborne lidar-retrieved canopy height in the WRF model

NASA Astrophysics Data System (ADS)

Lee, Junhong; Hong, Jinkyu

2016-06-01

Canopy height is closely related to biomass and aerodynamic properties, which regulate turbulent transfer of energy and mass at the soil-vegetation-atmosphere continuum. However, this key information has been prescribed as a constant value in a fixed plant functional type in atmospheric models. This paper is the first to report impacts of using realistic forest canopy height, retrieved from spaceborne lidar, on regional climate simulation by using the canopy height data in the Weather Research and Forecasting (WRF) model's land surface model. Numerical simulations were conducted over the Amazon Basin during summer season. Over this region, the lidar-retrieved canopy heights were higher than the default values used in the WRF, which are dependent only on plant functional type. By modifying roughness length and zero-plane displacement height, the change of canopy height resulted in changes in surface energy balance by regulating aerodynamic conductances and vertical temperature gradient, thus modifying the lifting condensation level and equivalent potential temperature in the atmospheric boundary layer. Our analysis also showed that the WRF model better reproduced the observed precipitation when lidar-retrieved canopy height was used over the Amazon Basin.

Preferred Negative Geotactic Orientation in Mobile Cells: Tetrahymena Results

NASA Technical Reports Server (NTRS)

Noever, David A.; Cronise, Raymond; Matsos, Helen C.

1994-01-01

For the protozoan species Tetrahymena a series of airplane experiments are reported, which varied gravity as an active laboratory parameter and tested for corresponding changes in geotaxic orientation of single cells. The airplane achieved altemating periods of low (0.01 g) and high (1.8 g, g = 980 cm/s) gravity by flying repeated Keplerian parabolas. The experimental design was undertaken to clearly distinguish gravity from competing aerodynamic and chemical gradients. In this way, each culture served as its own control, with gravity level alone determining the orientational changes. On average, 6.3% of the Tetrahymena oriented vertically in low gravity, while 27% oriented vertically in high-gravity phases. Simplified physical models are explored for describing these cell trajectores as a function of gravity, aerodynamic drag, and lift. The notable effect of gravity on turning behavior is emphasized as the biophysical cause of the observed negative geotaxis in Tetrahymena. A fundamental investigation of the biological gravity receptor (it it exists) and improved modeling for vertical migration in important types of ocean plankton motivate the present research.

Modeling State-Space Aeroelastic Systems Using a Simple Matrix Polynomial Approach for the Unsteady Aerodynamics

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.

Improved Skin Friction Interferometer

NASA Technical Reports Server (NTRS)

Westphal, R. V.; Bachalo, W. D.; Houser, M. H.

1986-01-01

An improved system for measuring aerodynamic skin friction which uses a dual-laser-beam oil-film interferometer was developed. Improvements in the optical hardware provided equal signal characteristics for each beam and reduced the cost and complexity of the system by replacing polarization rotation by a mirrored prism for separation of the two signals. An automated, objective, data-reduction procedure was implemented to eliminate tedious manual manipulation of the interferometry data records. The present system was intended for use in two-dimensional, incompressible flows over a smooth, level surface without pressure gradient, but the improvements discussed are not limited to this application.

Effective Thermal Conductivity of High Temperature Insulations for Reusable Launch Vehicles

NASA Technical Reports Server (NTRS)

Daryabeigi, Kamran

1999-01-01

An experimental apparatus was designed to measure the effective thermal conductivity of various high temperature insulations subject to large temperature gradients representative of typical launch vehicle re-entry aerodynamic heating conditions. The insulation sample cold side was maintained around room temperature, while the hot side was heated to temperatures as high as 1800 degrees Fahrenheit. The environmental pressure was varied from 0.0001 to 760 torr. All the measurements were performed in a dry gaseous nitrogen environment. The effective thermal conductivity of Saffil, Q-Fiber felt, Cerachrome, and three multi-layer insulation configurations were measured.

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, velocity, and acceleration sensors. This research demonstrates the feasibility of obtaining induced drag and lift forces through the use of distributed sensor technology with measured strain data. An active induced drag control system thus can be designed using the two computed aerodynamic forces, induced drag and lift, to improve the fuel efficiency of an aircraft. Interpolation elements between structural finite element grids and the CFD grids and centroids are successfully incorporated with the unsteady aeroelastic computation scheme. The most critical technology for the success of the proposed approach is the robust on-line parameter estimator, since the least-squares curve fitting method depends heavily on aeroelastic system frequencies and damping factors.

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.

Time-averaged aerodynamic loads on the vane sets of the 40- by 80-foot and 80- by 120-foot wind tunnel complex

NASA Technical Reports Server (NTRS)

Aoyagi, Kiyoshi; Olson, Lawrence E.; Peterson, Randall L.; Yamauchi, Gloria K.; Ross, James C.; Norman, Thomas R.

1987-01-01

Time-averaged aerodynamic loads are estimated for each of the vane sets in the National Full-Scale Aerodynamic Complex (NFAC). The methods used to compute global and local loads are presented. Experimental inputs used to calculate these loads are based primarily on data obtained from tests conducted in the NFAC 1/10-Scale Vane-Set Test Facility and from tests conducted in the NFAC 1/50-Scale Facility. For those vane sets located directly downstream of either the 40- by 80-ft test section or the 80- by 120-ft test section, aerodynamic loads caused by the impingement of model-generated wake vortices and model-generated jet and propeller wakes are also estimated.

Aerodynamic Classification of Swept-Wing Ice Accretion

NASA Technical Reports Server (NTRS)

Broeren, Andy; Diebold, Jeff; Bragg, Mike

2013-01-01

The continued design, certification and safe operation of swept-wing airplanes in icing conditions rely on the advancement of computational and experimental simulation methods for higher fidelity results over an increasing range of aircraft configurations and performance, and icing conditions. The current state-of-the-art in icing aerodynamics is mainly built upon a comprehensive understanding of two-dimensional geometries that does not currently exist for fundamentally three-dimensional geometries such as swept wings. The purpose of this report is to describe what is known of iced-swept-wing aerodynamics and to identify the type of research that is required to improve the current understanding. Following the method used in a previous review of iced-airfoil aerodynamics, this report proposes a classification of swept-wing ice accretion into four groups based upon unique flowfield attributes. These four groups are: ice roughness, horn ice, streamwise ice, and spanwise-ridge ice. For all of the proposed ice-shape classifications, relatively little is known about the three-dimensional flowfield and even less about the effect of Reynolds number and Mach number on these flowfields. The classifications and supporting data presented in this report can serve as a starting point as new research explores swept-wing aerodynamics with ice shapes. As further results are available, it is expected that these classifications will need to be updated and revised.

Cascade flutter analysis with transient response aerodynamics

NASA Technical Reports Server (NTRS)

Bakhle, Milind A.; Mahajan, Aparajit J.; Keith, Theo G., Jr.; Stefko, George L.

1991-01-01

Two methods for calculating linear frequency domain aerodynamic coefficients from a time marching Full Potential cascade solver are developed and verified. In the first method, the Influence Coefficient, solutions to elemental problems are superposed to obtain the solutions for a cascade in which all blades are vibrating with a constant interblade phase angle. The elemental problem consists of a single blade in the cascade oscillating while the other blades remain stationary. In the second method, the Pulse Response, the response to the transient motion of a blade is used to calculate influence coefficients. This is done by calculating the Fourier Transforms of the blade motion and the response. Both methods are validated by comparison with the Harmonic Oscillation method and give accurate results. The aerodynamic coefficients obtained from these methods are used for frequency domain flutter calculations involving a typical section blade structural model. An eigenvalue problem is solved for each interblade phase angle mode and the eigenvalues are used to determine aeroelastic stability. Flutter calculations are performed for two examples over a range of subsonic Mach numbers.

The effect of vocal fold vertical stiffness gradient on sound production

NASA Astrophysics Data System (ADS)

Geng, Biao; Xue, Qian; Zheng, Xudong

2015-11-01

It is observed in some experimental studies on canine vocal folds (VFs) that the inferior aspect of the vocal fold (VF) is much stiffer than the superior aspect under relatively large strain. Such vertical difference is supposed to promote the convergent-divergent shape during VF vibration and consequently facilitate the production of sound. In this study, we investigate the effect of vertical variation of VF stiffness on sound production using a numerical model. The vertical variation of stiffness is produced by linearly increasing the Young's modulus and shear modulus from the superior to inferior aspects in the cover layer, and its effect on phonation is examined in terms of aerodynamic and acoustic quantities such as flow rate, open quotient, skewness of flow wave form, sound intensity and vocal efficiency. The flow-induced vibration of the VF is solved with a finite element solver coupled with 1D Bernoulli equation, which is further coupled with a digital waveguide model. This study is designed to find out whether it's beneficial to artificially induce the vertical stiffness gradient by certain implanting material in VF restoring surgery, and if it is beneficial, what gradient is the most favorable.

An Investigation of the Aerodynamics and Cooling of a Horizontally-Opposed Engine Installation

NASA Technical Reports Server (NTRS)

Miley, S. J.

1977-01-01

A research program to investigate the aerodynamics of reciprocating aircraft engine cooling installations is discussed. Current results from a flight test program are presented concerning installation flow measurement methods. The influence of different inlet designs on installation cooling effectiveness and efficiency are described.

Program user's manual for an unsteady helicopter rotor-fuselage aerodynamic analysis

NASA Technical Reports Server (NTRS)

Lorber, Peter F.

1988-01-01

The Rotor-Fuselage Analysis is a method of calculating the aerodynamic reaction between a helicopter rotor and fuselage. This manual describes the structure and operation of the computer programs that make up the Rotor-Fuselage Analysis, programs which prepare the input and programs which display the output.

Engineering Aerodynamics

DTIC Science & Technology

1986-12-01

effective Reynolds Number5 to include the effet Of turbulence, which was supported in a convincing manner by the same ratio of 2.4 betwveen the Reynolds...iLIFT DEVICIS 143 methods incorl)orating various forms of flas) are shown on Figure 59. The other two methods, Boundary Layer Control and the Magnus ...Class Airship Hlull with Varying Lengths of Cylindric Midships," N.A.CA. Technical Report No. 138 (1922). 276 ENGINEERING AERODYNAMICS [Ch. 9 -- - - 2.0

The use of photogrammetric and stereophotogrammetric methods in aerodynamic experiments

NASA Astrophysics Data System (ADS)

Shmyreva, V. N.; Iakovlev, V. A.

The possibilities afforded by photogrammetry and stereophotogrammetry in current aerodynamic experiments, methods of image recording, and observation data processing are briefly reviewed. Some specific experiments illustrating the application of stereophotogrammetry are described. The applications discussed include the monitoring of model position in wind tunnels, determination of model deformations and displacements, determination of the deformations of real structural elements in static strength tests, and solution of a variety of problems in hydrodynamics.

Aerodynamic design and analysis system for supersonic aircraft. Part 1: General description and theoretical development

NASA Technical Reports Server (NTRS)

Middleton, W. D.; Lundry, J. L.

1975-01-01

An integrated system of computer programs has been developed for the design and analysis of supersonic configurations. The system uses linearized theory methods for the calculation of surface pressures and supersonic area rule concepts in combination with linearized theory for calculation of aerodynamic force coefficients. Interactive graphics are optional at the user's request. This part presents a general description of the system and describes the theoretical methods used.

Transonic Flow Computations Using Nonlinear Potential Methods

NASA Technical Reports Server (NTRS)

Holst, Terry L.; Kwak, Dochan (Technical Monitor)

2000-01-01

This presentation describes the state of transonic flow simulation using nonlinear potential methods for external aerodynamic applications. The presentation begins with a review of the various potential equation forms (with emphasis on the full potential equation) and includes a discussion of pertinent mathematical characteristics and all derivation assumptions. Impact of the derivation assumptions on simulation accuracy, especially with respect to shock wave capture, is discussed. Key characteristics of all numerical algorithm types used for solving nonlinear potential equations, including steady, unsteady, space marching, and design methods, are described. Both spatial discretization and iteration scheme characteristics are examined. Numerical results for various aerodynamic applications are included throughout the presentation to highlight key discussion points. The presentation ends with concluding remarks and recommendations for future work. Overall. nonlinear potential solvers are efficient, highly developed and routinely used in the aerodynamic design environment for cruise conditions. Published by Elsevier Science Ltd. All rights reserved.

Fitting aerodynamic forces in the Laplace domain: An application of a nonlinear nongradient technique to multilevel constrained optimization

NASA Technical Reports Server (NTRS)

Tiffany, S. H.; Adams, W. M., Jr.

1984-01-01

A technique which employs both linear and nonlinear methods in a multilevel optimization structure to best approximate generalized unsteady aerodynamic forces for arbitrary motion is described. Optimum selection of free parameters is made in a rational function approximation of the aerodynamic forces in the Laplace domain such that a best fit is obtained, in a least squares sense, to tabular data for purely oscillatory motion. The multilevel structure and the corresponding formulation of the objective models are presented which separate the reduction of the fit error into linear and nonlinear problems, thus enabling the use of linear methods where practical. Certain equality and inequality constraints that may be imposed are identified; a brief description of the nongradient, nonlinear optimizer which is used is given; and results which illustrate application of the method are presented.

Design and Analysis Tool for External-Compression Supersonic Inlets

NASA Technical Reports Server (NTRS)

Slater, John W.

2012-01-01

A computational tool named SUPIN has been developed to design and analyze external-compression supersonic inlets for aircraft at cruise speeds from Mach 1.6 to 2.0. The inlet types available include the axisymmetric outward-turning, two-dimensional single-duct, two-dimensional bifurcated-duct, and streamline-traced Busemann inlets. The aerodynamic performance is characterized by the flow rates, total pressure recovery, and drag. The inlet flowfield is divided into parts to provide a framework for the geometry and aerodynamic modeling and the parts are defined in terms of geometric factors. The low-fidelity aerodynamic analysis and design methods are based on analytic, empirical, and numerical methods which provide for quick analysis. SUPIN provides inlet geometry in the form of coordinates and surface grids useable by grid generation methods for higher-fidelity computational fluid dynamics (CFD) analysis. SUPIN is demonstrated through a series of design studies and CFD analyses were performed to verify some of the analysis results.

Multi-objective Optimization Strategies Using Adjoint Method and Game Theory in Aerodynamics

NASA Astrophysics Data System (ADS)

Tang, Zhili

2006-08-01

There are currently three different game strategies originated in economics: (1) Cooperative games (Pareto front), (2) Competitive games (Nash game) and (3) Hierarchical games (Stackelberg game). Each game achieves different equilibria with different performance, and their players play different roles in the games. Here, we introduced game concept into aerodynamic design, and combined it with adjoint method to solve multi-criteria aerodynamic optimization problems. The performance distinction of the equilibria of these three game strategies was investigated by numerical experiments. We computed Pareto front, Nash and Stackelberg equilibria of the same optimization problem with two conflicting and hierarchical targets under different parameterizations by using the deterministic optimization method. The numerical results show clearly that all the equilibria solutions are inferior to the Pareto front. Non-dominated Pareto front solutions are obtained, however the CPU cost to capture a set of solutions makes the Pareto front an expensive tool to the designer.

Tribological study of an aerodynamic thrust bearing in the supersonic regime

NASA Astrophysics Data System (ADS)

Dupuy, F.; Bou-Saïd, B.; Garcia, M.

2017-02-01

Nowadays, aerodynamic air thrust bearing are mainly used over a large panel of turbo-machineries. These systems become increasingly faster and up to operate in supersonic regime. They have not reached a sufficient level of research in terms of high speed. The possibility of an aerodynamic thrust bearing operating in a supersonic regime is studied. First, the air film dynamic study for high Reynolds number is based on the “modified Reynolds” equations, which take into account the inertia terms, the viscosity’s variation in the film thickness, and the turbulence. It’s an extension of the traditional model used in lubrication called the generalized Reynolds equation. The results show that a depression occur at the location of the change of slope of the tapper flat geometry. The hypothesis of presence of shock or rarefaction waves shows that the pressure gradient in the film thickness may be no longer negligible. The modified Reynolds equation may be not enough to describe the problem. A new system is built to study these phenomena: the Navier-Stokes equation are adapted to the film’s geometry. The dynamic air film’s behavior study in supersonic regime requires a shock capturing scheme called WENO scheme (“Weighted Essentially Non Oscillatory”), mainly used in shock and turbulence studies. The numerical results give the film behavior modelling of a fixed air thrust bearing pad. The evolution of the quantities shows that shock wave can occur in a thin film.

Estimation of effective aerodynamic roughness with altimeter measurements

NASA Technical Reports Server (NTRS)

Menenti, M.; Ritchie, J. C.

1992-01-01

A new method is presented for estimating the aerodynamic roughness length of heterogeneous land surfaces and complex landscapes using elevation measurements performed with an airborne laser altimeter and the Seasat radar altimeter. Land surface structure is characterized at increasing length scales by considering three basic landscape elements: (1) partial to complete canopies of herbaceous vegetation; (2) sparse obstacles (e.g., shrubs and trees); and (3) local relief. Measured parameters of land surface geometry are combined to obtain an effective aerodynamic roughness length which parameterizes the total atmosphere-land surface stress.

Determination of aerodynamic sensitivity coefficients based on the three-dimensional full potential equation

NASA Technical Reports Server (NTRS)

Elbanna, Hesham M.; Carlson, Leland A.

1992-01-01

The quasi-analytical approach is applied to the three-dimensional full potential equation to compute wing aerodynamic sensitivity coefficients in the transonic regime. Symbolic manipulation is used to reduce the effort associated with obtaining the sensitivity equations, and the large sensitivity system is solved using 'state of the art' routines. Results are compared to those obtained by the direct finite difference approach and both methods are evaluated to determine their computational accuracy and efficiency. The quasi-analytical approach is shown to be accurate and efficient for large aerodynamic systems.

Bayesian inference of nonlinear unsteady aerodynamics from aeroelastic limit cycle oscillations

NASA Astrophysics Data System (ADS)

Sandhu, Rimple; Poirel, Dominique; Pettit, Chris; Khalil, Mohammad; Sarkar, Abhijit

2016-07-01

A Bayesian model selection and parameter estimation algorithm is applied to investigate the influence of nonlinear and unsteady aerodynamic loads on the limit cycle oscillation (LCO) of a pitching airfoil in the transitional Reynolds number regime. At small angles of attack, laminar boundary layer trailing edge separation causes negative aerodynamic damping leading to the LCO. The fluid-structure interaction of the rigid, but elastically mounted, airfoil and nonlinear unsteady aerodynamics is represented by two coupled nonlinear stochastic ordinary differential equations containing uncertain parameters and model approximation errors. Several plausible aerodynamic models with increasing complexity are proposed to describe the aeroelastic system leading to LCO. The likelihood in the posterior parameter probability density function (pdf) is available semi-analytically using the extended Kalman filter for the state estimation of the coupled nonlinear structural and unsteady aerodynamic model. The posterior parameter pdf is sampled using a parallel and adaptive Markov Chain Monte Carlo (MCMC) algorithm. The posterior probability of each model is estimated using the Chib-Jeliazkov method that directly uses the posterior MCMC samples for evidence (marginal likelihood) computation. The Bayesian algorithm is validated through a numerical study and then applied to model the nonlinear unsteady aerodynamic loads using wind-tunnel test data at various Reynolds numbers.

Bayesian inference of nonlinear unsteady aerodynamics from aeroelastic limit cycle oscillations

DOE Office of Scientific and Technical Information (OSTI.GOV)

Sandhu, Rimple; Poirel, Dominique; Pettit, Chris

2016-07-01

A Bayesian model selection and parameter estimation algorithm is applied to investigate the influence of nonlinear and unsteady aerodynamic loads on the limit cycle oscillation (LCO) of a pitching airfoil in the transitional Reynolds number regime. At small angles of attack, laminar boundary layer trailing edge separation causes negative aerodynamic damping leading to the LCO. The fluid–structure interaction of the rigid, but elastically mounted, airfoil and nonlinear unsteady aerodynamics is represented by two coupled nonlinear stochastic ordinary differential equations containing uncertain parameters and model approximation errors. Several plausible aerodynamic models with increasing complexity are proposed to describe the aeroelastic systemmore » leading to LCO. The likelihood in the posterior parameter probability density function (pdf) is available semi-analytically using the extended Kalman filter for the state estimation of the coupled nonlinear structural and unsteady aerodynamic model. The posterior parameter pdf is sampled using a parallel and adaptive Markov Chain Monte Carlo (MCMC) algorithm. The posterior probability of each model is estimated using the Chib–Jeliazkov method that directly uses the posterior MCMC samples for evidence (marginal likelihood) computation. The Bayesian algorithm is validated through a numerical study and then applied to model the nonlinear unsteady aerodynamic loads using wind-tunnel test data at various Reynolds numbers.« less

Uncertainty quantification-based robust aerodynamic optimization of laminar flow nacelle

NASA Astrophysics Data System (ADS)

Xiong, Neng; Tao, Yang; Liu, Zhiyong; Lin, Jun

2018-05-01

The aerodynamic performance of laminar flow nacelle is highly sensitive to uncertain working conditions, especially the surface roughness. An efficient robust aerodynamic optimization method on the basis of non-deterministic computational fluid dynamic (CFD) simulation and Efficient Global Optimization (EGO)algorithm was employed. A non-intrusive polynomial chaos method is used in conjunction with an existing well-verified CFD module to quantify the uncertainty propagation in the flow field. This paper investigates the roughness modeling behavior with the γ-Ret shear stress transport model including modeling flow transition and surface roughness effects. The roughness effects are modeled to simulate sand grain roughness. A Class-Shape Transformation-based parametrical description of the nacelle contour as part of an automatic design evaluation process is presented. A Design-of-Experiments (DoE) was performed and surrogate model by Kriging method was built. The new design nacelle process demonstrates that significant improvements of both mean and variance of the efficiency are achieved and the proposed method can be applied to laminar flow nacelle design successfully.

Aerostructural Level Set Topology Optimization for a Common Research Model Wing

NASA Technical Reports Server (NTRS)

Dunning, Peter D.; Stanford, Bret K.; Kim, H. Alicia

2014-01-01

The purpose of this work is to use level set topology optimization to improve the design of a representative wing box structure for the NASA common research model. The objective is to minimize the total compliance of the structure under aerodynamic and body force loading, where the aerodynamic loading is coupled to the structural deformation. A taxi bump case was also considered, where only body force loads were applied. The trim condition that aerodynamic lift must balance the total weight of the aircraft is enforced by allowing the root angle of attack to change. The level set optimization method is implemented on an unstructured three-dimensional grid, so that the method can optimize a wing box with arbitrary geometry. Fast matching and upwind schemes are developed for an unstructured grid, which make the level set method robust and efficient. The adjoint method is used to obtain the coupled shape sensitivities required to perform aerostructural optimization of the wing box structure.

Characterization of the Space Shuttle Ascent Debris using CFD Methods

NASA Technical Reports Server (NTRS)

Murman, Scott M.; Aftosmis, Michael J.; Rogers, Stuart E.

2005-01-01

After video analysis of space shuttle flight STS-107's ascent showed that an object shed from the bipod-ramp region impacted the left wing, a transport analysis was initiated to determine a credible flight path and impact velocity for the piece of debris. This debris transport analysis was performed both during orbit, and after the subsequent re-entry accident. The analysis provided an accurate prediction of the velocity a large piece of foam bipod ramp would have as it impacted the wing leading edge. This prediction was corroborated by video analysis and fully-coupled CFD/six degree of freedom (DOF) simulations. While the prediction of impact velocity was accurate enough to predict critical damage in this case, one of the recommendations of the Columbia Accident Investigation Board (CAIB) for return-to-flight (RTF) was to analyze the complete debris environment experienced by the shuttle stack on ascent. This includes categorizing all possible debris sources, their probable geometric and aerodynamic characteristics, and their potential for damage. This paper is chiefly concerned with predicting the aerodynamic characteristics of a variety of potential debris sources (insulating foam and cork, nose-cone ablator, ice, ...) for the shuttle ascent configuration using CFD methods. These aerodynamic characteristics are used in the debris transport analysis to predict flight path, impact velocity and angle, and provide statistical variation to perform risk analyses where appropriate. The debris aerodynamic characteristics are difficult to determine using traditional methods, such as static or dynamic test data, due to the scaling requirements of simulating a typical debris event. The use of CFD methods has been a critical element for building confidence in the accuracy of the debris transport code by bridging the gap between existing aerodynamic data and the dynamics of full-scale, in-flight events.

Prediction of jump phenomena in rotationally-coupled maneuvers of aircraft, including nonlinear aerodynamic effects

NASA Technical Reports Server (NTRS)

Young, J. W.; Schy, A. A.; Johnson, K. G.

1977-01-01

An analytical method has been developed for predicting critical control inputs for which nonlinear rotational coupling may cause sudden jumps in aircraft response. The analysis includes the effect of aerodynamics which are nonlinear in angle of attack. The method involves the simultaneous solution of two polynomials in roll rate, whose coefficients are functions of angle of attack and the control inputs. Results obtained using this procedure are compared with calculated time histories to verify the validity of the method for predicting jump-like instabilities.

Numerical method of carbon-based material ablation effects on aero-heating for half-sphere

NASA Astrophysics Data System (ADS)

Wang, Jiang-Feng; Li, Jia-Wei; Zhao, Fa-Ming; Fan, Xiao-Feng

2018-05-01

A numerical method of aerodynamic heating with material thermal ablation effects for hypersonic half-sphere is presented. A surface material ablation model is provided to analyze the ablation effects on aero-thermal properties and structural heat conduction for thermal protection system (TPS) of hypersonic vehicles. To demonstrate its capability, applications for thermal analysis of hypersonic vehicles using carbonaceous ceramic ablators are performed and discussed. The numerical results show the high efficiency and validation of the method developed in thermal characteristics analysis of hypersonic aerodynamic heating.

Calculation of the distributed loads on the blades of individual multiblade propellers in axial flow using linear and nonlinear lifting surface theories

NASA Technical Reports Server (NTRS)

Pesetskaya, N. N.; Timofeev, I. YA.; Shipilov, S. D.

1988-01-01

In recent years much attention has been given to the development of methods and programs for the calculation of the aerodynamic characteristics of multiblade, saber-shaped air propellers. Most existing methods are based on the theory of lifting lines. Elsewhere, the theory of a lifting surface is used to calculate screw and lifting propellers. In this work, methods of discrete eddies are described for the calculation of the aerodynamic characteristics of propellers using the linear and nonlinear theories of lifting surfaces.

A rotor-aerodynamics-based wind estimation method using a quadrotor

NASA Astrophysics Data System (ADS)

Song, Yao; Luo, Bing; Meng, Qing-Hao

2018-02-01

Attempts to estimate horizontal wind by the quadrotor are reviewed. Wind estimations are realized by utilizing the quadrotor’s thrust change, which is caused by the wind’s effect on the rotors. The basis of the wind estimation method is the aerodynamic formula for the rotor’s thrust, which is verified and calibrated by experiments. A hardware-in-the-loop simulation (HILS) system was built as a testbed; its dynamic model and control structure are demonstrated. Verification experiments on the HILS system proved that the wind estimation method was effective.

Parameter Studies, time-dependent simulations and design with automated Cartesian methods

NASA Technical Reports Server (NTRS)

Aftosmis, Michael

2005-01-01

Over the past decade, NASA has made a substantial investment in developing adaptive Cartesian grid methods for aerodynamic simulation. Cartesian-based methods played a key role in both the Space Shuttle Accident Investigation and in NASA's return to flight activities. The talk will provide an overview of recent technological developments focusing on the generation of large-scale aerodynamic databases, automated CAD-based design, and time-dependent simulations with of bodies in relative motion. Automation, scalability and robustness underly all of these applications and research in each of these topics will be presented.

Computational Fluid Dynamics of Whole-Body Aircraft

NASA Astrophysics Data System (ADS)

Agarwal, Ramesh

1999-01-01

The current state of the art in computational aerodynamics for whole-body aircraft flowfield simulations is described. Recent advances in geometry modeling, surface and volume grid generation, and flow simulation algorithms have led to accurate flowfield predictions for increasingly complex and realistic configurations. As a result, computational aerodynamics has emerged as a crucial enabling technology for the design and development of flight vehicles. Examples illustrating the current capability for the prediction of transport and fighter aircraft flowfields are presented. Unfortunately, accurate modeling of turbulence remains a major difficulty in the analysis of viscosity-dominated flows. In the future, inverse design methods, multidisciplinary design optimization methods, artificial intelligence technology, and massively parallel computer technology will be incorporated into computational aerodynamics, opening up greater opportunities for improved product design at substantially reduced costs.

Real-Time Adaptive Least-Squares Drag Minimization for Performance Adaptive Aeroelastic Wing

NASA Technical Reports Server (NTRS)

Ferrier, Yvonne L.; Nguyen, Nhan T.; Ting, Eric

2016-01-01

This paper contains a simulation study of a real-time adaptive least-squares drag minimization algorithm for an aeroelastic model of a flexible wing aircraft. The aircraft model is based on the NASA Generic Transport Model (GTM). The wing structures incorporate a novel aerodynamic control surface known as the Variable Camber Continuous Trailing Edge Flap (VCCTEF). The drag minimization algorithm uses the Newton-Raphson method to find the optimal VCCTEF deflections for minimum drag in the context of an altitude-hold flight control mode at cruise conditions. The aerodynamic coefficient parameters used in this optimization method are identified in real-time using Recursive Least Squares (RLS). The results demonstrate the potential of the VCCTEF to improve aerodynamic efficiency for drag minimization for transport aircraft.

Development of V/STOL methodology based on a higher order panel method

NASA Technical Reports Server (NTRS)

Bhateley, I. C.; Howell, G. A.; Mann, H. W.

1983-01-01

The development of a computational technique to predict the complex flowfields of V/STOL aircraft was initiated in which a number of modules and a potential flow aerodynamic code were combined in a comprehensive computer program. The modules were developed in a building-block approach to assist the user in preparing the geometric input and to compute parameters needed to simulate certain flow phenomena that cannot be handled directly within a potential flow code. The PAN AIR aerodynamic code, which is higher order panel method, forms the nucleus of this program. PAN AIR's extensive capability for allowing generalized boundary conditions allows the modules to interact with the aerodynamic code through the input and output files, thereby requiring no changes to the basic code and easy replacement of updated modules.

Optimization of aerodynamic form of projectile for solving the problem of shooting range increasing

NASA Astrophysics Data System (ADS)

Lipanov, Alexey M.; Korolev, Stanislav A.; Rusyak, Ivan G.

2017-10-01

The article is devoted to the development of methods for solving the problem of external ballistics using a more complete system of motion equation taken into account the rotation and oscillation about the mass center and using aerodynamic coefficients of forces and moments which are calculated on the basis of modeling the hydrodynamics of flow around the projectile. Developed methods allows to study the basic ways of increasing the shooting range or artillery.

Calculation of the aerodynamic loading of swept and unswept flexible wings of arbitrary stiffness

NASA Technical Reports Server (NTRS)

Diederich, Franklin W

1950-01-01

A method is presented for calculating the aerodynamic loading, the divergence speed, and certain stability derivatives of swept and unswept wings and tail surfaces of arbitrary stiffness. Provision is made for using either stiffness curves and root rotation constants or structural influence coefficients in the analysis. Computing forms, tables of numerical constants required in the analysis, and an illustrative example are included to facilitate calculations by means of the method.

Colors Of Liquid Crystals Used To Measure Surface Shear Stresses

NASA Technical Reports Server (NTRS)

Reda, D. C.; Muratore, J. J., Jr.

1996-01-01

Developmental method of mapping shear stresses on aerodynamic surfaces involves observation, at multiple viewing angles, of colors of liquid-crystal surface coats illuminated by white light. Report describing method referenced in "Liquid Crystals Indicate Directions Of Surface Shear Stresses" (ARC-13379). Resulting maps of surface shear stresses contain valuable data on magnitudes and directions of skin friction forces associated with surface flows; data used to refine mathematical models of aerodynamics for research and design purposes.

Simulating the Refractive Index Structure Constant ({C}_{n}^{2}) in the Surface Layer at Antarctica with a Mesoscale Model

NASA Astrophysics Data System (ADS)

Qing, Chun; Wu, Xiaoqing; Li, Xuebin; Tian, Qiguo; Liu, Dong; Rao, Ruizhong; Zhu, Wenyue

2018-01-01

In this paper, we introduce an approach wherein the Weather Research and Forecasting (WRF) model is coupled with the bulk aerodynamic method to estimate the surface layer refractive index structure constant (C n 2) above Taishan Station in Antarctica. First, we use the measured meteorological parameters to estimate C n 2 using the bulk aerodynamic method, and second, we use the WRF model output parameters to estimate C n 2 using the bulk aerodynamic method. Finally, the corresponding C n 2 values from the micro-thermometer are compared with the C n 2 values estimated using the WRF model coupled with the bulk aerodynamic method. We analyzed the statistical operators—the bias, root mean square error (RMSE), bias-corrected RMSE (σ), and correlation coefficient (R xy )—in a 20 day data set to assess how this approach performs. In addition, we employ contingency tables to investigate the estimation quality of this approach, which provides complementary key information with respect to the bias, RMSE, σ, and R xy . The quantitative results are encouraging and permit us to confirm the fine performance of this approach. The main conclusions of this study tell us that this approach provides a positive impact on optimizing the observing time in astronomical applications and provides complementary key information for potential astronomical sites.

A dynamic aerodynamic resistance approach to calculate high resolution sensible heat fluxes in urban areas

NASA Astrophysics Data System (ADS)

Crawford, Ben; Grimmond, Sue; Kent, Christoph; Gabey, Andrew; Ward, Helen; Sun, Ting; Morrison, William

2017-04-01

Remotely sensed data from satellites have potential to enable high-resolution, automated calculation of urban surface energy balance terms and inform decisions about urban adaptations to environmental change. However, aerodynamic resistance methods to estimate sensible heat flux (QH) in cities using satellite-derived observations of surface temperature are difficult in part due to spatial and temporal variability of the thermal aerodynamic resistance term (rah). In this work, we extend an empirical function to estimate rah using observational data from several cities with a broad range of surface vegetation land cover properties. We then use this function to calculate spatially and temporally variable rah in London based on high-resolution (100 m) land cover datasets and in situ meteorological observations. In order to calculate high-resolution QH based on satellite-observed land surface temperatures, we also develop and employ novel methods to i) apply source area-weighted averaging of surface and meteorological variables across the study spatial domain, ii) calculate spatially variable, high-resolution meteorological variables (wind speed, friction velocity, and Obukhov length), iii) incorporate spatially interpolated urban air temperatures from a distributed sensor network, and iv) apply a modified Monte Carlo approach to assess uncertainties with our results, methods, and input variables. Modeled QH using the aerodynamic resistance method is then compared to in situ observations in central London from a unique network of scintillometers and eddy-covariance measurements.

An initial investigation into methods of computing transonic aerodynamic sensitivity coefficients

NASA Technical Reports Server (NTRS)

Carlson, Leland A.

1994-01-01

The primary accomplishments of the project are as follows: (1) Using the transonic small perturbation equation as a flowfield model, the project demonstrated that the quasi-analytical method could be used to obtain aerodynamic sensitivity coefficients for airfoils at subsonic, transonic, and supersonic conditions for design variables such as Mach number, airfoil thickness, maximum camber, angle of attack, and location of maximum camber. It was established that the quasi-analytical approach was an accurate method for obtaining aerodynamic sensitivity derivatives for airfoils at transonic conditions and usually more efficient than the finite difference approach. (2) The usage of symbolic manipulation software to determine the appropriate expressions and computer coding associated with the quasi-analytical method for sensitivity derivatives was investigated. Using the three dimensional fully conservative full potential flowfield model, it was determined that symbolic manipulation along with a chain rule approach was extremely useful in developing a combined flowfield and quasi-analytical sensitivity derivative code capable of considering a large number of realistic design variables. (3) Using the three dimensional fully conservative full potential flowfield model, the quasi-analytical method was applied to swept wings (i.e. three dimensional) at transonic flow conditions. (4) The incremental iterative technique has been applied to the three dimensional transonic nonlinear small perturbation flowfield formulation, an equivalent plate deflection model, and the associated aerodynamic and structural discipline sensitivity equations; and coupled aeroelastic results for an aspect ratio three wing in transonic flow have been obtained.

Comparison of aerodynamic coefficients obtained from theoretical calculations wind tunnel tests and flight tests data reduction for the alpha jet aircraft

NASA Technical Reports Server (NTRS)

Guiot, R.; Wunnenberg, H.

1980-01-01

The methods by which aerodynamic coefficients are determined and discussed. These include: calculations, wind tunnel experiments and experiments in flight for various prototypes of the Alpha Jet. A comparison of obtained results shows good correlation between expectations and in-flight test results.

Ventricular-Fold Dynamics in Human Phonation

ERIC Educational Resources Information Center

Bailly, Lucie; Bernardoni, Nathalie Henrich; Müller, Frank; Rohlfs, Anna-Katharina; Hess, Markus

2014-01-01

Purpose: In this study, the authors aimed (a) to provide a classification of the ventricular-fold dynamics during voicing, (b) to study the aerodynamic impact of these motions on vocal-fold vibrations, and (c) to assess whether ventricular-fold oscillations could be sustained by aerodynamic coupling with the vocal folds. Method: A 72-sample…

Aerodynamic and Nonlinear Dynamic Acoustic Analysis of Tension Asymmetry in Excised Canine Larynges

ERIC Educational Resources Information Center

Devine, Erin E.; Bulleit, Erin E.; Hoffman, Matthew R.; McCulloch, Timothy M.; Jiang, Jack J.

2012-01-01

Purpose: To model tension asymmetry caused by superior laryngeal nerve paralysis (SLNP) in excised larynges and apply perturbation, nonlinear dynamic, and aerodynamic analyses. Method: SLNP was modeled in 8 excised larynges using sutures and weights to mimic cricothyroid (CT) muscle function. Weights were removed from one side to create tension…

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.

SUPIN: A Computational Tool for Supersonic Inlet Design

NASA Technical Reports Server (NTRS)

Slater, John W.

2016-01-01

A computational tool named SUPIN is being developed to design and analyze the aerodynamic performance of supersonic inlets. The inlet types available include the axisymmetric pitot, three-dimensional pitot, axisymmetric outward-turning, two-dimensional single-duct, two-dimensional bifurcated-duct, and streamline-traced inlets. The aerodynamic performance is characterized by the flow rates, total pressure recovery, and drag. The inlet flow-field is divided into parts to provide a framework for the geometry and aerodynamic modeling. Each part of the inlet is defined in terms of geometric factors. The low-fidelity aerodynamic analysis and design methods are based on analytic, empirical, and numerical methods which provide for quick design and analysis. SUPIN provides inlet geometry in the form of coordinates, surface angles, and cross-sectional areas. SUPIN can generate inlet surface grids and three-dimensional, structured volume grids for use with higher-fidelity computational fluid dynamics (CFD) analysis. Capabilities highlighted in this paper include the design and analysis of streamline-traced external-compression inlets, modeling of porous bleed, and the design and analysis of mixed-compression inlets. CFD analyses are used to verify the SUPIN results.

Design of off-statistics axial-flow fans by means of vortex law optimization

NASA Astrophysics Data System (ADS)

Lazari, Andrea; Cattanei, Andrea

2014-12-01

Off-statistics input data sets are common in axial-flow fans design and may easily result in some violation of the requirements of a good aerodynamic blade design. In order to circumvent this problem, in the present paper, a solution to the radial equilibrium equation is found which minimizes the outlet kinetic energy and fulfills the aerodynamic constraints, thus ensuring that the resulting blade has acceptable aerodynamic performance. The presented method is based on the optimization of a three-parameters vortex law and of the meridional channel size. The aerodynamic quantities to be employed as constraints are individuated and their suitable ranges of variation are proposed. The method is validated by means of a design with critical input data values and CFD analysis. Then, by means of systematic computations with different input data sets, some correlations and charts are obtained which are analogous to classic correlations based on statistical investigations on existing machines. Such new correlations help size a fan of given characteristics as well as study the feasibility of a given design.

The design and development of transonic multistage compressors

NASA Technical Reports Server (NTRS)

Ball, C. L.; Steinke, R. J.; Newman, F. A.

1988-01-01

The development of the transonic multistage compressor is reviewed. Changing trends in design and performance parameters are noted. These changes are related to advances in compressor aerodynamics, computational fluid mechanics and other enabling technologies. The parameters normally given to the designer and those that need to be established during the design process are identified. Criteria and procedures used in the selection of these parameters are presented. The selection of tip speed, aerodynamic loading, flowpath geometry, incidence and deviation angles, blade/vane geometry, blade/vane solidity, stage reaction, aerodynamic blockage, inlet flow per unit annulus area, stage/overall velocity ratio, and aerodynamic losses are considered. Trends in these parameters both spanwise and axially through the machine are highlighted. The effects of flow mixing and methods for accounting for the mixing in the design process are discussed.

Aerodynamic loads on buses due to crosswind gusts: extended analysis

NASA Astrophysics Data System (ADS)

Drugge, Lars; Juhlin, Magnus

2010-12-01

The objective of this work is to use inverse simulations on measured vehicle data in order to estimate the aerodynamic loads on a bus when exposed to crosswind situations. Tyre forces, driver input, wind velocity and vehicle response were measured on a typical coach when subjected to natural crosswind gusts. Based on these measurements and a detailed MBS vehicle model, the aerodynamic loads were estimated through inverse simulations. In order to estimate the lift force, roll and pitch moments in addition to the lateral force and yaw moment, the simulation model was extended by also incorporating the estimation of the vertical road disturbances. The proposed method enables the estimation of aerodynamic loads due to crosswind gusts without using a full scale wind tunnel adapted for crosswind excitation.

Aerodynamic study of time-trial helmets in cycling racing using CFD analysis.

PubMed

Beaumont, F; Taiar, R; Polidori, G; Trenchard, H; Grappe, F

2018-01-23

The aerodynamic drag of three different time-trial cycling helmets was analyzed numerically for two different cyclist head positions. Computational Fluid Dynamics (CFD) methods were used to investigate the detailed airflow patterns around the cyclist for a constant velocity of 15 m/s without wind. The CFD simulations have focused on the aerodynamic drag effects in terms of wall shear stress maps and pressure coefficient distributions on the cyclist/helmet system. For a given head position, the helmet shape, by itself, obtained a weak effect on a cyclist's aerodynamic performance (<1.5%). However, by varying head position, a cyclist significantly influences aerodynamic performance; the maximum difference between both positions being about 6.4%. CFD results have also shown that both helmet shape and head position significantly influence drag forces, pressure and wall shear stress distributions on the whole cyclist's body due to the change in the near-wake behavior and in location of corresponding separation and attachment areas around the cyclist. Copyright © 2017 Elsevier Ltd. All rights reserved.

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.

Individual Monitoring of Vocal Effort With Relative Fundamental Frequency: Relationships With Aerodynamics and Listener Perception.

PubMed

Lien, Yu-An S; Michener, Carolyn M; Eadie, Tanya L; Stepp, Cara E

2015-06-01

The acoustic measure relative fundamental frequency (RFF) was investigated as a potential objective measure to track variations in vocal effort within and across individuals. Twelve speakers with healthy voices created purposeful modulations in their vocal effort during speech tasks. RFF and an aerodynamic measure of vocal effort, the ratio of sound pressure level to subglottal pressure level, were estimated from the aerodynamic and acoustic signals. Twelve listeners also judged the speech samples for vocal effort using the visual sort and rate method. Relationships between RFF and both the aerodynamic and perceptual measures of vocal effort were weak across speakers (R2 = .06-.26). Within speakers, relationships were variable but much stronger on average (R2 = .45-.56). RFF showed stronger relationships between both the aerodynamic and perceptual measures of vocal effort when examined within individuals versus across individuals. Future work is necessary to establish these relationships in individuals with voice disorders across the therapeutic process.

2D Inviscid and Viscous Inverse Design Using Continuous Adjoint and Lax-Wendroff Formulation

NASA Astrophysics Data System (ADS)

Proctor, Camron Lisle

The continuous adjoint (CA) technique for optimization and/or inverse-design of aerodynamic components has seen nearly 30 years of documented success in academia. The benefits of using CA versus a direct sensitivity analysis are shown repeatedly in the literature. However, the use of CA in industry is relatively unheard-of. The sparseness of industry contributions to the field may be attributed to the tediousness of the derivation and/or to the difficulties in implementation due to the lack of well-documented adjoint numerical methods. The focus of this work has been to thoroughly document the techniques required to build a two-dimensional CA inverse-design tool. To this end, this work begins with a short background on computational fluid dynamics (CFD) and the use of optimization tools in conjunction with CFD tools to solve aerodynamic optimization problems. A thorough derivation of the continuous adjoint equations and the accompanying gradient calculations for inviscid and viscous constraining equations follows the introduction. Next, the numerical techniques used for solving the partial differential equations (PDEs) governing the flow equations and the adjoint equations are described. Numerical techniques for the supplementary equations are discussed briefly. Subsequently, a verification of the efficacy of the inverse design tool, for the inviscid adjoint equations as well as possible numerical implementation pitfalls are discussed. The NACA0012 airfoil is used as an initial airfoil and surface pressure distribution and the NACA16009 is used as the desired pressure and vice versa. Using a Savitsky-Golay gradient filter, convergence (defined as a cost function<1E-5) is reached in approximately 220 design iteration using 121 design variables. The inverse-design using inviscid adjoint equations results are followed by the discussion of the viscous inverse design results and techniques used to further the convergence of the optimizer. The relationship between limiting step-size and convergence in a line-search optimization is shown to slightly decrease the final cost function at significant computational cost. A gradient damping technique is presented and shown to increase the convergence rate for the optimization in viscous problems, at a negligible increase in computational cost, but is insufficient to converge the solution. Systematically including adjacent surface vertices in the perturbation of a design variable, also a surface vertex, is shown to affect the convergence capability of the viscous optimizer. Finally, a comparison of using inviscid adjoint equations, as opposed to viscous adjoint equations, on viscous flow is presented, and the inviscid adjoint paired with viscous flow is found to reduce the cost function further than the viscous adjoint for the presented problem.

Optical surface pressure measurements: Accuracy and application field evaluation

NASA Astrophysics Data System (ADS)

Bukov, A.; Mosharov, V.; Orlov, A.; Pesetsky, V.; Radchenko, V.; Phonov, S.; Matyash, S.; Kuzmin, M.; Sadovskii, N.

1994-07-01

Optical pressure measurement (OPM) is a new pressure measurement method rapidly developed in several aerodynamic research centers: TsAGI (Russia), Boeing, NASA, McDonnell Douglas (all USA), and DLR (Germany). Present level of OPM-method provides its practice as standard experimental method of aerodynamic investigations in definite application fields. Applications of OPM-method are determined mainly by its accuracy. The accuracy of OPM-method is determined by the errors of three following groups: (1) errors of the luminescent pressure sensor (LPS) itself, such as uncompensated temperature influence, photo degradation, temperature and pressure hysteresis, variation of the LPS parameters from point to point on the model surface, etc.; (2) errors of the measurement system, such as noise of the photodetector, nonlinearity and nonuniformity of the photodetector, time and temperature offsets, etc.; and (3) methodological errors, owing to displacement and deformation of the model in an airflow, a contamination of the model surface, scattering of the excitation and luminescent light from the model surface and test section walls, etc. OPM-method allows getting total error of measured pressure not less than 1 percent. This accuracy is enough to visualize the pressure field and allows determining total and distributed aerodynamic loads and solving some problems of local aerodynamic investigations at transonic and supersonic velocities. OPM is less effective at low subsonic velocities (M less than 0.4), and for precise measurements, for example, an airfoil optimization. Current limitations of the OPM-method are discussed on an example of the surface pressure measurements and calculations of the integral loads on the wings of canard-aircraft model. The pressure measurement system and data reduction methods used on these tests are also described.

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.

Spectral methods and their implementation to solution of aerodynamic and fluid mechanic problems

NASA Technical Reports Server (NTRS)

Streett, C. L.

1987-01-01

Fundamental concepts underlying spectral collocation methods, especially pertaining to their use in the solution of partial differential equations, are outlined. Theoretical accuracy results are reviewed and compared with results from test problems. A number of practical aspects of the construction and use of spectral methods are detailed, along with several solution schemes which have found utility in applications of spectral methods to practical problems. Results from a few of the successful applications of spectral methods to problems of aerodynamic and fluid mechanic interest are then outlined, followed by a discussion of the problem areas in spectral methods and the current research under way to overcome these difficulties.

DOE Office of Scientific and Technical Information (OSTI.GOV)

Mendenhall, M.R.

The present volume discusses tactical missile aerodynamic drag, drag-prediction methods for axisymmetric missile bodies, an aerodynamic heating analysis for supersonic missiles, a component buildup method for engineering analysis of missiles at low-to-high angles of attack, experimental and analytical methods for missiles with noncircular fuselages, and a vortex-cloud model for body vortex shedding and tracking. Also discussed are panel methods with vorticity effects and corrections for nonlinear compressibility, supersonic full-potential methods for missile body analysis, space-marching Euler solvers, the time-asymptotic Euler/Navier-Stokes methods for subsonic and transonic flows, 3D boundary layers on missiles, Navier-Stokes analyses of flows over slender airframes, and themore » interaction of exhaust plumes with missile airframes.« less

An Efficient Multiblock Method for Aerodynamic Analysis and Design on Distributed Memory Systems

NASA Technical Reports Server (NTRS)

Reuther, James; Alonso, Juan Jose; Vassberg, John C.; Jameson, Antony; Martinelli, Luigi

1997-01-01

The work presented in this paper describes the application of a multiblock gridding strategy to the solution of aerodynamic design optimization problems involving complex configurations. The design process is parallelized using the MPI (Message Passing Interface) Standard such that it can be efficiently run on a variety of distributed memory systems ranging from traditional parallel computers to networks of workstations. Substantial improvements to the parallel performance of the baseline method are presented, with particular attention to their impact on the scalability of the program as a function of the mesh size. Drag minimization calculations at a fixed coefficient of lift are presented for a business jet configuration that includes the wing, body, pylon, aft-mounted nacelle, and vertical and horizontal tails. An aerodynamic design optimization is performed with both the Euler and Reynolds Averaged Navier-Stokes (RANS) equations governing the flow solution and the results are compared. These sample calculations establish the feasibility of efficient aerodynamic optimization of complete aircraft configurations using the RANS equations as the flow model. There still exists, however, the need for detailed studies of the importance of a true viscous adjoint method which holds the promise of tackling the minimization of not only the wave and induced components of drag, but also the viscous drag.

Field sampling of loose erodible material: A new method to consider the full particle-size range

NASA Astrophysics Data System (ADS)

Klose, Martina; Gill, Thomas E.

2017-04-01

The aerodynamic entrainment of sand and dust is determined by the atmospheric forces exerted onto the soil surface and by the soil-surface condition. If aerodynamic forces are strong enough to generate sand and dust lifting, the entrained sediment amount still critically depends on the supply of loose particles readily available for lifting. This loose erodible material (LEM) is sometimes defined as the thin layer of loose particles on top of a crusted surface. Here, we more generally define LEM as loose particles or particle aggregates available for entrainment, which may or may not overlay a soil crust. Field sampling of LEM is difficult and only few attempts have been made. Motivated by saltation as the most efficient process to generate dust emission, methods have focused on capturing LEM in the sand-size range or on determining the potential of a soil surface to be eroded by aerodynamic forces and particle impacts. Here, our focus is to capture the full particle-size distribution of LEM in situ, including the dust and sand-size range, to investigate the potential and likelihood of dust emission mechanisms (aerodynamic entrainment, saltation bombardment, aggregate disintegration) to occur. A new vacuum method is introduced and its capability to sample LEM without significant alteration of the LEM particle-size distribution is investigated.

Research on the drag reduction performance induced by the counterflowing jet for waverider with variable blunt radii

NASA Astrophysics Data System (ADS)

Li, Shi-bin; Wang, Zhen-guo; Barakos, George N.; Huang, Wei; Steijl, Rene

2016-10-01

Waverider will endure the huge aero-heating in the hypersonic flow, thus, it need be blunt for the leading edge. However, the aerodynamic performance will decrease for the blunt waverider because of the drag hoik. How to improve the aerodynamic performance and reduce the drag and aero-heating is very important. The variable blunt radii method will improve the aerodynamic performance, however, the huge aero-heating and bow shock wave at the head is still serious. In the current study, opposing jet is used in the waverider with variable blunt radii to improve its performance. The three-dimensional coupled implicit Reynolds-averaged Navier-Stokes(RANS) equation and the two equation SST k-ω turbulence model have been utilized to obtain the flow field properties. The numerical method has been validated against the available experimental data in the open literature. The obtained results show that the L/D will drop 7-8% when R changes from 2 to 8. The lift coefficient will increase, and the drag coefficient almost keeps the same when the variable blunt radii method is adopted, and the L/D will increase. The variable blunt radii method is very useful to improve the whole characteristics of blunt waverider and the L/D can improve 3%. The combination of the variable blunt radii method and opposing jet is a novel way to improve the whole performance of blunt waverider, and L/D can improve 4-5%. The aperture as a novel way of opposing jet is suitable for blunt waverider and also useful to improve the aerodynamic and aerothermodynamic characteristics of waverider in the hypersonic flow. There is the optimal P0in/P0 that can make the detached shock wave reattach the lower surface again so that the blunt waverider can get the better aerodynamic performance.

A fourth order Euler/Navier-Stokes prediction method for the aerodynamics and aeroelasticity of hovering rotor blades

NASA Astrophysics Data System (ADS)

Smith, Marilyn Jones

Some of the computational issues relating to the development of a three-dimensional fourth-order compact Euler/Navier-Stokes methodology for rotary wing flows and its coupling with an elastic rotor blade beam structural model have been explored. The compact Euler/NavierStokes method is used to predict the aerodynamic loads on an isolated rotor blade. Because the scheme is fourth-order, fewer grid nodes are necessary to predict loads with the same accuracy as traditional second order methodologies on finer grids. Grid and numerical parameter optimizations were performed to examine the changes in the predictive capabilities of the higher-order scheme. Comparisons were made with experimental data for a rotor using NACA 0012 airfoil sections and a rectangular planform with no twist. Simulations for both lifting and non-lifting configurations at various tip Mach numbers were performed. This Euler/Navier-Stokes methodology can be applied to rotor blades with either rigid-blade or elastic-beam-structural models to determine the steady-state response in hovering flight. The blade is represented by a geometrically nonlinear beam model which accounts for coupled flap bending, lead-lag bending and torsion. Moderately large displacements and rotations due to structural deformations can be simulated. The analysis has been performed for blade configurations having uniform mass and stiffness, no twist, and no chordwise offsets of the elastic and tension axes, as well as the center of mass. The results are compared with a panel method coupled with the same structural dynamics model. Computations have been made to predict the aerodynamic deflections for the rotor in hover. A starting solution using initial deflections predicted by aeroelastic analyses with a two-dimensional aerodynamic model was investigated. The present Euler/Navier-Stokes method using a momentum wake and a contracting vortex wake shows the impact on the aeroelastic deflections of a three-dimensional aerodynamic module which includes rotational and viscous effects, particularly at higher collective pitch angles. The differences in the aeroelastic predictions using fully coupled and loosely coupled aerodynamic analyses are examined. The induced wake plays a critical role in determining the final equilibrium tip deflections.

Experimental investigation of turbine disk cavity aerodynamics and heat transfer

NASA Technical Reports Server (NTRS)

Daniels, W. A.; Johnson, B. V.

1993-01-01

An experimental investigation of turbine disk cavity aerodynamics and heat transfer was conducted to provide an experimental data base that can guide the aerodynamic and thermal design of turbine disks and blade attachments for flow conditions and geometries simulating those of the space shuttle main engine (SSME) turbopump drive turbines. Experiments were conducted to define the nature of the aerodynamics and heat transfer of the flow within the disk cavities and blade attachments of a large scale model simulating the SSME turbopump drive turbines. These experiments include flow between the main gas path and the disk cavities, flow within the disk cavities, and leakage flows through the blade attachments and labyrinth seals. Air was used to simulate the combustion products in the gas path. Air and carbon dioxide were used to simulate the coolants injected at three locations in the disk cavities. Trace amounts of carbon dioxide were used to determine the source of the gas at selected locations on the rotors, the cavity walls, and the interstage seal. The measurements on the rotor and stationary walls in the forward and aft cavities showed that the coolant effectiveness was 90 percent or greater when the coolant flow rate was greater than the local free disk entrainment flow rate and when room temperature air was used as both coolant and gas path fluid. When a coolant-to-gas-path density ratio of 1.51 was used in the aft cavity, the coolant effectiveness on the rotor was also 90 percent or greater at the aforementioned condition. However, the coolant concentration on the stationary wall was 60 to 80 percent at the aforementioned condition indicating a more rapid mixing of the coolant and flow through the rotor shank passages. This increased mixing rate was attributed to the destabilizing effects of the adverse density gradients.

Spacecraft flight control system design selection process for a geostationary communication satellite

NASA Technical Reports Server (NTRS)

Barret, C.

1992-01-01

The Earth's first artificial satellite, Sputnik 1, slowly tumbled in orbit. The first U.S. satellite, Explorer 1, also tumbled out of control. Now, as we launch the Mars observer and the Cassini spacecraft, stability and control have become higher priorities. The flight control system design selection process is reviewed using as an example a geostationary communication satellite which is to have a life expectancy of 10 to 14 years. Disturbance torques including aerodynamic, magnetic, gravity gradient, solar, micrometeorite, debris, collision, and internal torques are assessed to quantify the disturbance environment so that the required compensating torque can be determined. Then control torque options, including passive versus active, momentum control, bias momentum, spin stabilization, dual spin, gravity gradient, magnetic, reaction wheels, control moment gyros, nutation dampers, inertia augmentation techniques, three-axis control, reactions control system (RCS), and RCS sizing, are considered. A flight control system design is then selected and preliminary stability criteria are met by the control gains selection.

PyFly: A fast, portable aerodynamics simulator

DOE Office of Scientific and Technical Information (OSTI.GOV)

Garcia, Daniel; Ghommem, M.; Collier, Nathaniel O.

Here, we present a fast, user-friendly implementation of a potential flow solver based on the unsteady vortex lattice method (UVLM), namely PyFly. UVLM computes the aerodynamic loads applied on lifting surfaces while capturing the unsteady effects such as the added mass forces, the growth of bound circulation, and the wake while assuming that the flow separation location is known a priori. This method is based on discretizing the body surface into a lattice of vortex rings and relies on the Biot–Savart law to construct the velocity field at every point in the simulated domain. We introduce the pointwise approximation approachmore » to simulate the interactions of the far-field vortices to overcome the computational burden associated with the classical implementation of UVLM. The computational framework uses the Python programming language to provide an easy to handle user interface while the computational kernels are written in Fortran. The mixed language approach enables high performance regarding solution time and great flexibility concerning easiness of code adaptation to different system configurations and applications. The computational tool predicts the unsteady aerodynamic behavior of multiple moving bodies (e.g., flapping wings, rotating blades, suspension bridges) subject to incoming air. The aerodynamic simulator can also deal with enclosure effects, multi-body interactions, and B-spline representation of body shapes. Finally, we simulate different aerodynamic problems to illustrate the usefulness and effectiveness of PyFly.« less

PyFly: A fast, portable aerodynamics simulator

DOE PAGES

Garcia, Daniel; Ghommem, M.; Collier, Nathaniel O.; ...

2018-03-14

Here, we present a fast, user-friendly implementation of a potential flow solver based on the unsteady vortex lattice method (UVLM), namely PyFly. UVLM computes the aerodynamic loads applied on lifting surfaces while capturing the unsteady effects such as the added mass forces, the growth of bound circulation, and the wake while assuming that the flow separation location is known a priori. This method is based on discretizing the body surface into a lattice of vortex rings and relies on the Biot–Savart law to construct the velocity field at every point in the simulated domain. We introduce the pointwise approximation approachmore » to simulate the interactions of the far-field vortices to overcome the computational burden associated with the classical implementation of UVLM. The computational framework uses the Python programming language to provide an easy to handle user interface while the computational kernels are written in Fortran. The mixed language approach enables high performance regarding solution time and great flexibility concerning easiness of code adaptation to different system configurations and applications. The computational tool predicts the unsteady aerodynamic behavior of multiple moving bodies (e.g., flapping wings, rotating blades, suspension bridges) subject to incoming air. The aerodynamic simulator can also deal with enclosure effects, multi-body interactions, and B-spline representation of body shapes. Finally, we simulate different aerodynamic problems to illustrate the usefulness and effectiveness of PyFly.« less

Approach to Modeling Boundary Layer Ingestion Using a Fully Coupled Propulsion-RANS Model

NASA Technical Reports Server (NTRS)

Gray, Justin S.; Mader, Charles A.; Kenway, Gaetan K. W.; Martins, Joaquim R. R. A.

2017-01-01

Airframe-propulsion integration concepts that use boundary layer ingestion have the potential to reduce aircraft fuel burn. One concept that has been recently explored is NASA's Starc-ABL aircraft configuration, which offers the potential for 12% mission fuel burn reduction by using a turbo-electric propulsion system with an aft-mounted electrically driven boundary layer ingestion propulsor. This large potential for improved performance motivates a more detailed study of the boundary layer ingestion propulsor design, but to date, analyses of boundary layer ingestion have used uncoupled methods. These methods account for only aerodynamic effects on the propulsion system or propulsion system effects on the aerodynamics, but not both simultaneously. This work presents a new approach for building fully coupled propulsive-aerodynamic models of boundary layer ingestion propulsion systems. A 1D thermodynamic cycle analysis is coupled to a RANS simulation to model the Starc-ABL aft propulsor at a cruise condition and the effects variation in propulsor design on performance are examined. The results indicates that both propulsion and aerodynamic effects contribute equally toward the overall performance and that the fully coupled model yields substantially different results compared to uncoupled. The most significant finding is that boundary layer ingestion, while offering substantial fuel burn savings, introduces throttle dependent aerodynamics effects that need to be accounted for. This work represents a first step toward the multidisciplinary design optimization of boundary layer ingestion propulsion systems.

Study of aerodynamic technology for single-cruise engine V/STOL fighter/attack aircraft

NASA Technical Reports Server (NTRS)

Driggers, H. H.; Powers, S. A.; Roush, R. T.

1982-01-01

A conceptual design analysis is performed on a single engine V/STOL supersonic fighter/attack concept powered by a series flow tandem fan propulsion system. Forward and aft mounted fans have independent flow paths for V/STOL operation and series flow in high speed flight. Mission, combat and V/STOL performance is calculated. Detailed aerodynamic estimates are made and aerodynamic uncertainties associated with the configuration and estimation methods identified. A wind tunnel research program is developed to resolve principal uncertainties and establish a data base for the baseline configuration and parametric variations.

AGARD standard aeroelastic configurations for dynamic response. Candidate configuration I.-wing 445.6

NASA Technical Reports Server (NTRS)

Yates, E. Carson, Jr.

1987-01-01

To promote the evaluation of existing and emerging unsteady aerodynamic codes and methods for applying them to aeroelastic problems, especially for the transonic range, a limited number of aerodynamic configurations and experimental dynamic response data sets are to be designated by the AGARD Structures and Materials Panel as standards for comparison. This set is a sequel to that established several years ago for comparisons of calculated and measured aerodynamic pressures and forces. This report presents the information needed to perform flutter calculations for the first candidate standard configuration for dynamic response along with the related experimental flutter data.

Research on inverse, hybrid and optimization problems in engineering sciences with emphasis on turbomachine aerodynamics: Review of Chinese advances

NASA Technical Reports Server (NTRS)

Liu, Gao-Lian

1991-01-01

Advances in inverse design and optimization theory in engineering fields in China are presented. Two original approaches, the image-space approach and the variational approach, are discussed in terms of turbomachine aerodynamic inverse design. Other areas of research in turbomachine aerodynamic inverse design include the improved mean-streamline (stream surface) method and optimization theory based on optimal control. Among the additional engineering fields discussed are the following: the inverse problem of heat conduction, free-surface flow, variational cogeneration of optimal grid and flow field, and optimal meshing theory of gears.

Simulation on a car interior aerodynamic noise control based on statistical energy analysis

NASA Astrophysics Data System (ADS)

Chen, Xin; Wang, Dengfeng; Ma, Zhengdong

2012-09-01

How to simulate interior aerodynamic noise accurately is an important question of a car interior noise reduction. The unsteady aerodynamic pressure on body surfaces is proved to be the key effect factor of car interior aerodynamic noise control in high frequency on high speed. In this paper, a detail statistical energy analysis (SEA) model is built. And the vibra-acoustic power inputs are loaded on the model for the valid result of car interior noise analysis. The model is the solid foundation for further optimization on car interior noise control. After the most sensitive subsystems for the power contribution to car interior noise are pointed by SEA comprehensive analysis, the sound pressure level of car interior aerodynamic noise can be reduced by improving their sound and damping characteristics. The further vehicle testing results show that it is available to improve the interior acoustic performance by using detailed SEA model, which comprised by more than 80 subsystems, with the unsteady aerodynamic pressure calculation on body surfaces and the materials improvement of sound/damping properties. It is able to acquire more than 2 dB reduction on the central frequency in the spectrum over 800 Hz. The proposed optimization method can be looked as a reference of car interior aerodynamic noise control by the detail SEA model integrated unsteady computational fluid dynamics (CFD) and sensitivity analysis of acoustic contribution.

TURNS - A free-wake Euler/Navier-Stokes numerical method for helicopter rotors

NASA Technical Reports Server (NTRS)

Srinivasan, G. R.; Baeder, J. D.

1993-01-01

Computational capabilities of a numerical procedure, called TURNS (transonic unsteady rotor Navier-Stokes), to calculate the aerodynamics and acoustics (high-speed impulsive noise) out to several rotor diameters are summarized. The procedure makes it possible to obtain the aerodynamics and acoustics information in one single calculation. The vortical wave and its influence, as well as the acoustics, are captured as part of the overall flowfield solution. The accuracy and suitability of the TURNS method is demonstrated through comparisons with experimental data.

Fluid mechanics of dynamic stall. II - Prediction of full scale characteristics

NASA Technical Reports Server (NTRS)

Ericsson, L. E.; Reding, J. P.

1988-01-01

Analytical extrapolations are made from experimental subscale dynamics to predict full scale characteristics of dynamic stall. The method proceeds by establishing analytic relationships between dynamic and static aerodynamic characteristics induced by viscous flow effects. The method is then validated by predicting dynamic test results on the basis of corresponding static test data obtained at the same subscale flow conditions, and the effect of Reynolds number on the static aerodynamic characteristics are determined from subscale to full scale flow conditions.

A collection of flow visualization techniques used in the Aerodynamic Research Branch

NASA Technical Reports Server (NTRS)

1984-01-01

Theoretical and experimental research on unsteady aerodynamic flows is discussed. Complex flow fields that involve separations, vortex interactions, and transonic flow effects were investigated. Flow visualization techniques are used to obtain a global picture of the flow phenomena before detailed quantitative studies are undertaken. A wide variety of methods are used to visualize fluid flow and a sampling of these methods is presented. It is emphasized that the visualization technique is a thorough quantitative analysis and subsequent physical understanding of these flow fields.

A method of calculation on the airloading of vertical axis wind turbine

NASA Astrophysics Data System (ADS)

Azuma, A.; Kimura, S.

A new method of analyzing the aerodynamic characteristics of the Darrieus Vertical-Axis Wind Turbine (VAWT) by applying the local circulation method is described. The validity of this method is confirmed by analyzing the air load acting on a curved blade. The azimuthwise variation of spanwise airloading, torque, and longitudinal forces are accurately calculated for a variety of operational conditions. The results are found to be in good agreement with experimental ones obtained elsewhere. It is concluded that the present approach can calculate the aerodynamic characteristics of the VAWT with much less computational time than that used by the free vortex model.

A Numerical Model of Unsteady, Subsonic Aeroelastic Behavior. Ph.D Thesis

NASA Technical Reports Server (NTRS)

Strganac, Thomas W.

1987-01-01

A method for predicting unsteady, subsonic aeroelastic responses was developed. The technique accounts for aerodynamic nonlinearities associated with angles of attack, vortex-dominated flow, static deformations, and unsteady behavior. The fluid and the wing together are treated as a single dynamical system, and the equations of motion for the structure and flow field are integrated simultaneously and interactively in the time domain. The method employs an iterative scheme based on a predictor-corrector technique. The aerodynamic loads are computed by the general unsteady vortex-lattice method and are determined simultaneously with the motion of the wing. Because the unsteady vortex-lattice method predicts the wake as part of the solution, the history of the motion is taken into account; hysteresis is predicted. Two models are used to demonstrate the technique: a rigid wing on an elastic support experiencing plunge and pitch about the elastic axis, and an elastic wing rigidly supported at the root chord experiencing spanwise bending and twisting. The method can be readily extended to account for structural nonlinearities and/or substitute aerodynamic load models. The time domain solution coupled with the unsteady vortex-lattice method provides the capability of graphically depicting wing and wake motion.

The Effect of Speaking Rate on Velopharyngeal Function in Healthy Speakers

ERIC Educational Resources Information Center

Gauster, Andrea; Yunusova, Yana; Zajac, David

2010-01-01

The purpose of this study was to assess the effect of speaking rate variation on aerodynamic and acoustic measures of velopharyngeal (VP) function. Twenty-seven healthy adult speakers (14 males, 13 females) participated in the study. The modified pressure-flow method was used to collect aerodynamic data of /m/ and /p/ segments in the word…

Individual Monitoring of Vocal Effort with Relative Fundamental Frequency: Relationships with Aerodynamics and Listener Perception

ERIC Educational Resources Information Center

Lien, Yu-An S.; Michener, Carolyn M.; Eadie, Tanya L.; Stepp, Cara E.

2015-01-01

Purpose: The acoustic measure relative fundamental frequency (RFF) was investigated as a potential objective measure to track variations in vocal effort within and across individuals. Method: Twelve speakers with healthy voices created purposeful modulations in their vocal effort during speech tasks. RFF and an aerodynamic measure of vocal effort,…

Journal of Aeronautics.

DTIC Science & Technology

1982-07-21

aerodynamic tool for design of elastic aircraft. Several numerical examples are given and some dynamical problems of elastic aircraft are also discussed...Qiangang, Wu Changlin, Jian Zheng Northwestern Polytechnical University Abstract: A numerical metbod,6* ted for predicting the aerodynamic characte- ristics... Numerical value calculation method is one important means of the present research on elastic aircraft pneumatic characteristics. Be- cause this

Unconventional tail configurations for transport aircraft

NASA Astrophysics Data System (ADS)

Sánchez-Carmona, A.; Cuerno-Rejado, C.; García-Hernández, L.

2017-06-01

This article presents the bases of a methodology in order to size unconventional tail configurations for transport aircraft. The case study of this paper is a V-tail con¦guration. Firstly, an aerodynamic study is developed for determining stability derivatives and aerodynamic forces. The objective is to size a tail such as it develops at least the same static stability derivatives than a conventional reference aircraft. The optimum is obtained minimizing its weight. The weight is estimated through two methods: adapted Farrar£s method and a statistical method. The solution reached is heavier than the reference, but it reduces the wetted area.

Structural dynamics and aerodynamics measurements of biologically inspired flexible flapping wings.

PubMed

Wu, P; Stanford, B K; Sällström, E; Ukeiley, L; Ifju, P G

2011-03-01

Flapping wing flight as seen in hummingbirds and insects poses an interesting unsteady aerodynamic problem: coupling of wing kinematics, structural dynamics and aerodynamics. There have been numerous studies on the kinematics and aerodynamics in both experimental and computational cases with both natural and artificial wings. These studies tend to ignore wing flexibility; however, observation in nature affirms that passive wing deformation is predominant and may be crucial to the aerodynamic performance. This paper presents a multidisciplinary experimental endeavor in correlating a flapping micro air vehicle wing's aeroelasticity and thrust production, by quantifying and comparing overall thrust, structural deformation and airflow of six pairs of hummingbird-shaped membrane wings of different properties. The results show that for a specific spatial distribution of flexibility, there is an effective frequency range in thrust production. The wing deformation at the thrust-productive frequencies indicates the importance of flexibility: both bending and twisting motion can interact with aerodynamic loads to enhance wing performance under certain conditions, such as the deformation phase and amplitude. By measuring structural deformations under the same aerodynamic conditions, beneficial effects of passive wing deformation can be observed from the visualized airflow and averaged thrust. The measurements and their presentation enable observation and understanding of the required structural properties for a thrust effective flapping wing. The intended passive responses of the different wings follow a particular pattern in correlation to their aerodynamic performance. Consequently, both the experimental technique and data analysis method can lead to further studies to determine the design principles for micro air vehicle flapping wings.

[Acoustic and aerodynamic characteristics of the oesophageal voice].

PubMed

Vázquez de la Iglesia, F; Fernández González, S

2005-12-01

The aim of the study is to determine the physiology and pathophisiology of esophageal voice according to objective aerodynamic and acoustic parameters (quantitative and qualitative parameters). Our subjects were comprised of 33 laryngectomized patients (all male) that underwent aerodynamic, acoustic and perceptual protocol. There is a statistical association between acoustic and aerodynamic qualitative parameters (phonation flow chart type, sound spectrum, perceptual analysis) among quantitative parameters (neoglotic pressure, phonation flow, phonation time, fundamental frequency, maximum intensity sound level, speech rate). Nevertheles, not always such observations bring practical resources to clinical practice. We consider that the facts studied may enable us to add, pragmatically, new resources to the more effective vocal rehabilitation to these patients. The physiology of esophageal voice is well understood by the method we have applied, also seeking for rehabilitation, improving oral communication skills in the laryngectomee population.

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.

Investigation on the forced response of a radial turbine under aerodynamic excitations

NASA Astrophysics Data System (ADS)

Ma, Chaochen; Huang, Zhi; Qi, Mingxu

2016-04-01

Rotor blades in a radial turbine with nozzle guide vanes typically experience harmonic aerodynamic excitations due to the rotor stator interaction. Dynamic stresses induced by the harmonic excitations can result in high cycle fatigue (HCF) of the blades. A reliable prediction method for forced response issue is essential to avoid the HCF problem. In this work, the forced response mechanisms were investigated based on a fluid structure interaction (FSI) method. Aerodynamic excitations were obtained by three-dimensional unsteady computational fluid dynamics (CFD) simulation with phase shifted periodic boundary conditions. The first two harmonic pressures were determined as the primary components of the excitation and applied to finite element (FE) model to conduct the computational structural dynamics (CSD) simulation. The computed results from the harmonic forced response analysis show good agreement with the predictions of Singh's advanced frequency evaluation (SAFE) diagram. Moreover, the mode superposition method used in FE simulation offers an efficient way to provide quantitative assessments of mode response levels and resonant strength.

Full-envelope aerodynamic modeling of the Harrier aircraft

NASA Technical Reports Server (NTRS)

Mcnally, B. David

1986-01-01

A project to identify a full-envelope model of the YAV-8B Harrier using flight-test and parameter identification techniques is described. As part of the research in advanced control and display concepts for V/STOL aircraft, a full-envelope aerodynamic model of the Harrier is identified, using mathematical model structures and parameter identification methods. A global-polynomial model structure is also used as a basis for the identification of the YAV-8B aerodynamic model. State estimation methods are used to ensure flight data consistency prior to parameter identification.Equation-error methods are used to identify model parameters. A fixed-base simulator is used extensively to develop flight test procedures and to validate parameter identification software. Using simple flight maneuvers, a simulated data set was created covering the YAV-8B flight envelope from about 0.3 to 0.7 Mach and about -5 to 15 deg angle of attack. A singular value decomposition implementation of the equation-error approach produced good parameter estimates based on this simulated data set.

Influence of a heated leading edge on boundary layer growth, stability, and transition

DOE Office of Scientific and Technical Information (OSTI.GOV)

Landrum, D.B.; Macha, J.M.

1987-06-01

This paper presents the results of a combined theoretical and experimental study of the influence of a heated leading edge on the growth, stability, and transition of a two-dimensional boundary layer. The findings are directly applicable to aircraft wings and nacelles that use surface heating for anti-icing protection. The potential effects of the non-adiabatic condition are particularly important for laminar-flow sections where even small perturbations can result in significantly degraded aerodynamic performance. The results of the study give new insight to the fundamental coupling between streamwise pressure gradient and surface heat flux in laminar and transitional boundary layers. 13 references.

Influence of a heated leading edge on boundary layer growth, stability, and transition

DOE Office of Scientific and Technical Information (OSTI.GOV)

Landrum, D.B.; Macha, J.M.

1987-01-01

This paper presents the results of a combined theoretical and experimental study of the influence of a heated leading edge on the growth, stability, and transition of a two-dimensional boundary layer. The findings are directly applicable to aircraft wings and nacelles that use surface heating for anti-icing protection. The potential effects of the non-adiabatic condition are particularly important for laminar-flow sections where even small perturbations can result in significantly degraded aerodynamic performance. The results of the study give new insight to the fundamental coupling between streamwise pressure gradient and surface heat flux in laminar and transitional boundary layers.

Wavy-Planform Helicopter Blades Make Less Noise

NASA Technical Reports Server (NTRS)

Brooks, Thomas F.

2004-01-01

Wavy-planform rotor blades for helicopters have been investigated for the first time in an effort to reduce noise. Two of the main sources of helicopter noise are blade/vortex interaction (BVI) and volume displacement. (The noise contributed by volume displacement is termed thickness noise.) The reduction in noise generated by a wavyplanform blade, relative to that generated by an otherwise equivalent straight-planform blade, affects both main sources: (1) the BVI noise is reduced through smoothing and defocusing of the aerodynamic loading on the blade and (2) the thickness noise is reduced by reducing gradients of thickness with respect to listeners on the ground.

A generic multi-flex-body dynamics, controls simulation tool for space station

NASA Technical Reports Server (NTRS)

London, Ken W.; Lee, John F.; Singh, Ramen P.; Schubele, Buddy

1991-01-01

An order (n) multiflex body Space Station simulation tool is introduced. The flex multibody modeling is generic enough to model all phases of Space Station from build up through to Assembly Complete configuration and beyond. Multibody subsystems such as the Mobile Servicing System (MSS) undergoing a prescribed translation and rotation are also allowed. The software includes aerodynamic, gravity gradient, and magnetic field models. User defined controllers can be discrete or continuous. Extensive preprocessing of 'body by body' NASTRAN flex data is built in. A significant aspect, too, is the integrated controls design capability which includes model reduction and analytic linearization.

Skylab viscous damper study

NASA Technical Reports Server (NTRS)

1978-01-01

The proposed magnetically anchored viscous fluid damper can maintain the Skylab in a gravity-gradient stabilized mode at the anticipated reboost altitudes. The parameters influencing damper performance (and thereby affecting the degree of risk) are: (1) amount of skylab pitch bias in the orbit plane which will result from aerodynamic trim conditions of the post-reboost configuration Skylab; (2) the lowest altitude to which the post-reboost Skylab will be allowed to decay prior to the next rendezvous; (3) maximum allowable weight and size of the proposed damper in order to match shuttle/TRS mission constraints; (4) the amount of magnetic materials expected to be in the vicinity of the damper.

Wind modeling and lateral control for automatic landing

NASA Technical Reports Server (NTRS)

Holley, W. E.; Bryson, A. E., Jr.

1975-01-01

For the purposes of aircraft control system design and analysis, the wind can be characterized by a mean component which varies with height and by turbulent components which are described by the von Karman correlation model. The aircraft aero-dynamic forces and moments depend linearly on uniform and gradient gust components obtained by averaging over the aircraft's length and span. The correlations of the averaged components are then approximated by the outputs of linear shaping filters forced by white noise. The resulting model of the crosswind shear and turbulence effects is used in the design of a lateral control system for the automatic landing of a DC-8 aircraft.

A review on design of experiments and surrogate models in aircraft real-time and many-query aerodynamic analyses

NASA Astrophysics Data System (ADS)

Yondo, Raul; Andrés, Esther; Valero, Eusebio

2018-01-01

Full scale aerodynamic wind tunnel testing, numerical simulation of high dimensional (full-order) aerodynamic models or flight testing are some of the fundamental but complex steps in the various design phases of recent civil transport aircrafts. Current aircraft aerodynamic designs have increase in complexity (multidisciplinary, multi-objective or multi-fidelity) and need to address the challenges posed by the nonlinearity of the objective functions and constraints, uncertainty quantification in aerodynamic problems or the restrained computational budgets. With the aim to reduce the computational burden and generate low-cost but accurate models that mimic those full order models at different values of the design variables, Recent progresses have witnessed the introduction, in real-time and many-query analyses, of surrogate-based approaches as rapid and cheaper to simulate models. In this paper, a comprehensive and state-of-the art survey on common surrogate modeling techniques and surrogate-based optimization methods is given, with an emphasis on models selection and validation, dimensionality reduction, sensitivity analyses, constraints handling or infill and stopping criteria. Benefits, drawbacks and comparative discussions in applying those methods are described. Furthermore, the paper familiarizes the readers with surrogate models that have been successfully applied to the general field of fluid dynamics, but not yet in the aerospace industry. Additionally, the review revisits the most popular sampling strategies used in conducting physical and simulation-based experiments in aircraft aerodynamic design. Attractive or smart designs infrequently used in the field and discussions on advanced sampling methodologies are presented, to give a glance on the various efficient possibilities to a priori sample the parameter space. Closing remarks foster on future perspectives, challenges and shortcomings associated with the use of surrogate models by aircraft industrial aerodynamicists, despite their increased interest among the research communities.

Reynolds number influences in aeronautics

NASA Technical Reports Server (NTRS)

Bushnell, Dennis M.; Yip, Long P.; Yao, Chung-Sheng; Lin, John C.; Lawing, Pierce L.; Batina, John T.; Hardin, Jay C.; Horvath, Thomas J.; Fenbert, James W.; Domack, Christopher S.

1993-01-01

Reynolds number, a measure of the ratio of inertia to viscous forces, is a fundamental similarity parameter for fluid flows and therefore, would be expected to have a major influence in aerodynamics and aeronautics. Reynolds number influences are generally large, but monatomic, for attached laminar (continuum) flow; however, laminar flows are easily separated, inducing even stronger, non-monatomic, Reynolds number sensitivities. Probably the strongest Reynolds number influences occur in connection with transitional flow behavior. Transition can take place over a tremendous Reynolds number range, from the order of 20 x 10(exp 3) for 2-D free shear layers up to the order of 100 x 10(exp 6) for hypersonic boundary layers. This variability in transition behavior is especially important for complex configurations where various vehicle and flow field elements can undergo transition at various Reynolds numbers, causing often surprising changes in aerodynamics characteristics over wide ranges in Reynolds number. This is further compounded by the vast parameterization associated with transition, in that any parameter which influences mean viscous flow development (e.g., pressure gradient, flow curvature, wall temperature, Mach number, sweep, roughness, flow chemistry, shock interactions, etc.), and incident disturbance fields (acoustics, vorticity, particulates, temperature spottiness, even electro static discharges) can alter transition locations to first order. The usual method of dealing with the transition problem is to trip the flow in the generally lower Reynolds number wind tunnel to simulate the flight turbulent behavior. However, this is not wholly satisfactory as it results in incorrectly scaled viscous region thicknesses and cannot be utilized at all for applications such as turbine blades and helicopter rotors, nacelles, leading edge and nose regions, and High Altitude Long Endurance and hypersonic airbreathers where the transitional flow is an innately critical portion of the problem.

Review of Research on Low-Profile Vortex Generators to Control Boundary-Layer Separation

NASA Technical Reports Server (NTRS)

Lin, John C.

2002-01-01

An in-depth review of boundary-layer flow-separation control by a passive method using low-profile vortex generators is presented. The generators are defined as those with a device height between 10% and 50% of the boundary layer thickness. Key results are presented for several research efforts, all of which were performed within the past decade and a half where the majority of these works emphasize experimentation with some recent efforts on numerical simulations. Topics of discussion consist of both basic fluid dynamics and applied aerodynamics research. The fluid dynamics research includes comparative studies on separation control effectiveness as well as device-induced vortex characterization and correlation. The comparative studies cover the controlling of low-speed separated flows in adverse pressure gradient and supersonic shock-induced separation. The aerodynamics research includes several applications for aircraft performance enhancement and covers a wide range of speeds. Significant performance improvements are achieved through increased lift and/or reduced drag for various airfoils-low-Reynolds number, high-lift, and transonic-as well as highly swept wings. Performance enhancements for non-airfoil applications include aircraft interior noise reduction, inlet flow distortion alleviation inside compact ducts, and a more efficient overwing fairing. The low-profile vortex generators are best for being applied to applications where flow-separation locations are relatively fixed and the generators can be placed reasonably close upstream of the separation. Using the approach of minimal near-wall proturbances through substantially reduced device height, these devices can produce streamwise vortices just strong enough to overcome the separation without unnecessarily persisting within the boundary layer once the flow-control objective is achieved. Practical advantages of low-profile vortex generators, such as their inherent simplicity and low device drag, are demonstrated to be critically important for many applications as well.

Convergence characteristics of nonlinear vortex-lattice methods for configuration aerodynamics

NASA Technical Reports Server (NTRS)

Seginer, A.; Rusak, Z.; Wasserstrom, E.

1983-01-01

Nonlinear panel methods have no proof for the existence and uniqueness of their solutions. The convergence characteristics of an iterative, nonlinear vortex-lattice method are, therefore, carefully investigated. The effects of several parameters, including (1) the surface-paneling method, (2) an integration method of the trajectories of the wake vortices, (3) vortex-grid refinement, and (4) the initial conditions for the first iteration on the computed aerodynamic coefficients and on the flow-field details are presented. The convergence of the iterative-solution procedure is usually rapid. The solution converges with grid refinement to a constant value, but the final value is not unique and varies with the wing surface-paneling and wake-discretization methods within some range in the vicinity of the experimental result.

Estimation of Unsteady Aerodynamics in the Wake of a Freely Flying European Starling (Sturnus vulgaris)

PubMed Central

Ben-Gida, Hadar; Kirchhefer, Adam; Taylor, Zachary J.; Bezner-Kerr, Wayne; Guglielmo, Christopher G.; Kopp, Gregory A.; Gurka, Roi

2013-01-01

Wing flapping is one of the most widespread propulsion methods found in nature; however, the current understanding of the aerodynamics in bird wakes is incomplete. The role of the unsteady motion in the flow and its contribution to the aerodynamics is still an open question. In the current study, the wake of a freely flying European starling has been investigated using long-duration high-speed Particle Image Velocimetry (PIV) in the near wake. Kinematic analysis of the wings and body of the bird has been performed using additional high-speed cameras that recorded the bird movement simultaneously with the PIV measurements. The wake evolution of four complete wingbeats has been characterized through reconstruction of the time-resolved data, and the aerodynamics in the wake have been analyzed in terms of the streamwise forces acting on the bird. The profile drag from classical aerodynamics was found to be positive during most of the wingbeat cycle, yet kinematic images show that the bird does not decelerate. It is shown that unsteady aerodynamics are necessary to satisfy the drag/thrust balance by approximating the unsteady drag term. These findings may shed light on the flight efficiency of birds by providing a partial answer to how they minimize drag during flapping flight. PMID:24278243

Estimation of unsteady aerodynamics in the wake of a freely flying European starling (Sturnus vulgaris).

PubMed

Ben-Gida, Hadar; Kirchhefer, Adam; Taylor, Zachary J; Bezner-Kerr, Wayne; Guglielmo, Christopher G; Kopp, Gregory A; Gurka, Roi

2013-01-01

Wing flapping is one of the most widespread propulsion methods found in nature; however, the current understanding of the aerodynamics in bird wakes is incomplete. The role of the unsteady motion in the flow and its contribution to the aerodynamics is still an open question. In the current study, the wake of a freely flying European starling has been investigated using long-duration high-speed Particle Image Velocimetry (PIV) in the near wake. Kinematic analysis of the wings and body of the bird has been performed using additional high-speed cameras that recorded the bird movement simultaneously with the PIV measurements. The wake evolution of four complete wingbeats has been characterized through reconstruction of the time-resolved data, and the aerodynamics in the wake have been analyzed in terms of the streamwise forces acting on the bird. The profile drag from classical aerodynamics was found to be positive during most of the wingbeat cycle, yet kinematic images show that the bird does not decelerate. It is shown that unsteady aerodynamics are necessary to satisfy the drag/thrust balance by approximating the unsteady drag term. These findings may shed light on the flight efficiency of birds by providing a partial answer to how they minimize drag during flapping flight.

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.

Aerodynamics of High-Lift Configuration Civil Aircraft Model in JAXA

NASA Astrophysics Data System (ADS)

Yokokawa, Yuzuru; Murayama, Mitsuhiro; Ito, Takeshi; Yamamoto, Kazuomi

This paper presents basic aerodynamics and stall characteristics of the high-lift configuration aircraft model JSM (JAXA Standard Model). During research process of developing high-lift system design method, wind tunnel testing at JAXA 6.5m by 5.5m low-speed wind tunnel and Navier-Stokes computation on unstructured hybrid mesh were performed for a realistic configuration aircraft model equipped with high-lift devices, fuselage, nacelle-pylon, slat tracks and Flap Track Fairings (FTF), which was assumed 100 passenger class modern commercial transport aircraft. The testing and the computation aimed to understand flow physics and then to obtain some guidelines for designing a high performance high-lift system. As a result of the testing, Reynolds number effects within linear region and stall region were observed. Analysis of static pressure distribution and flow visualization gave the knowledge to understand the aerodynamic performance. CFD could capture the whole characteristics of basic aerodynamics and clarify flow mechanism which governs stall characteristics even for complicated geometry and its flow field. This collaborative work between wind tunnel testing and CFD is advantageous for improving or has improved the aerodynamic performance.

New technology in turbine aerodynamics

NASA Technical Reports Server (NTRS)

Glassman, A. J.; Moffitt, T. P.

1972-01-01

A cursory review is presented of some of the recent work that has been done in turbine aerodynamic research at NASA-Lewis Research Center. Topics discussed include the aerodynamic effect of turbine coolant, high work-factor (ratio of stage work to square of blade speed) turbines, and computer methods for turbine design and performance prediction. An extensive bibliography is included. Experimental cooled-turbine aerodynamics programs using two-dimensional cascades, full annular cascades, and cold rotating turbine stage tests are discussed with some typical results presented. Analytically predicted results for cooled blade performance are compared to experimental results. The problems and some of the current programs associated with the use of very high work factors for fan-drive turbines of high-bypass-ratio engines are discussed. Turbines currently being investigated make use of advanced blading concepts designed to maintain high efficiency under conditions of high aerodynamic loading. Computer programs have been developed for turbine design-point performance, off-design performance, supersonic blade profile design, and the calculation of channel velocities for subsonic and transonic flow fields. The use of these programs for the design and analysis of axial and radial turbines is discussed.

Nasalance Scores of Children with Repaired Cleft Palate Who Exhibit Normal Velopharyngeal Closure during Aerodynamic Testing

ERIC Educational Resources Information Center

Zajac, David J.

2013-01-01

Purpose: To determine if children with repaired cleft palate and normal velopharyngeal (VP) closure as determined by aerodynamic testing exhibit greater acoustic nasalance than control children without cleft palate. Method: Pressure-flow procedures were used to identify 2 groups of children based on VP closure during the production of /p/ in the…

Aerodynamic and Acoustic Effects of Abrupt Frequency Changes in Excised Larynges

ERIC Educational Resources Information Center

Alipour, Fariborz; Finnegan, Eileen M.; Scherer, Ronald C.

2009-01-01

Purpose: To determine the aerodynamic and acoustic effects due to a sudden change from chest to falsetto register or vice versa. It was hypothesized that the continuous change in subglottal pressure and flow rate alone (pressure-flow sweep [PFS]) can trigger a mode change in the canine larynx. Method: Ten canine larynges were each mounted over a…

Transonic Dynamics Tunnel Force and Pressure Data Acquired on the HSR Rigid Semispan Model

NASA Technical Reports Server (NTRS)

Schuster, David M.; Rausch, Russ D.

1999-01-01

This report describes the aerodynamic data acquired on the High Speed Research Rigid Semispan Model (HSR-RSM) during NASA Langley Transonic Dynamics Tunnel (TDT) Test 520 conducted from 18 March to 4 April, 1996. The purpose of this test was to assess the aerodynamic character of a rigid high speed civil transport wing. The wing was fitted with a single trailing edge control surface which was both steadily deflected and oscillated during the test to investigate the response of the aerodynamic data to steady and unsteady control motion. Angle-of-attack and control surface deflection polars at subsonic, transonic and low-supersonic Mach numbers were obtained in the tunnel?s heavy gas configuration. Unsteady pressure and steady loads data were acquired on the wing, while steady pressures were measured on the fuselage. These data were reduced using a variety of methods, programs and computer systems. The reduced data was ultimately compiled onto a CD-ROM volume which was distributed to HSR industry team members in July, 1996. This report documents the methods used to acquire and reduce the data, and provides an assessment of the quality, repeatability, and overall character of the aerodynamic data measured during this test.

The Effect of Systematic Error in Forced Oscillation Testing

NASA Technical Reports Server (NTRS)

Williams, Brianne Y.; Landman, Drew; Flory, Isaac L., IV; Murphy, Patrick C.

2012-01-01

One of the fundamental problems in flight dynamics is the formulation of aerodynamic forces and moments acting on an aircraft in arbitrary motion. Classically, conventional stability derivatives are used for the representation of aerodynamic loads in the aircraft equations of motion. However, for modern aircraft with highly nonlinear and unsteady aerodynamic characteristics undergoing maneuvers at high angle of attack and/or angular rates the conventional stability derivative model is no longer valid. Attempts to formulate aerodynamic model equations with unsteady terms are based on several different wind tunnel techniques: for example, captive, wind tunnel single degree-of-freedom, and wind tunnel free-flying techniques. One of the most common techniques is forced oscillation testing. However, the forced oscillation testing method does not address the systematic and systematic correlation errors from the test apparatus that cause inconsistencies in the measured oscillatory stability derivatives. The primary objective of this study is to identify the possible sources and magnitude of systematic error in representative dynamic test apparatuses. Sensitivities of the longitudinal stability derivatives to systematic errors are computed, using a high fidelity simulation of a forced oscillation test rig, and assessed using both Design of Experiments and Monte Carlo methods.

The Effects of Crosswind Flight on Rotor Harmonic Noise Radiation

NASA Technical Reports Server (NTRS)

Greenwood, Eric; Sim, Ben W.

2013-01-01

In order to develop recommendations for procedures for helicopter source noise characterization, the effects of crosswinds on main rotor harmonic noise radiation are assessed using a model of the Bell 430 helicopter. Crosswinds are found to have a significant effect on Blade-Vortex Interaction (BVI) noise radiation when the helicopter is trimmed with the fuselage oriented along the inertial flight path. However, the magnitude of BVI noise remains unchanged when the pilot orients the fuselage along the aerodynamic velocity vector, crabbing for zero aerodynamic sideslip. The effects of wind gradients on BVI noise are also investigated and found to be smaller in the crosswind direction than in the headwind direction. The effects of crosswinds on lower harmonic noise sources at higher flight speeds are also assessed. In all cases, the directivity of radiated noise is somewhat changed by the crosswind. The model predictions agree well with flight test data for the Bell 430 helicopter captured under various wind conditions. The results of this investigation would suggest that flight paths for future acoustic flight testing are best aligned across the prevailing wind direction to minimize the effects of winds on noise measurements when wind cannot otherwise be avoided.

Modeling of wind-initiated liberation of fungal propagules from host plant leaves

NASA Astrophysics Data System (ADS)

Gonzalinajec, Trevor

2014-11-01

Successful airborne propagule dispersal must begin with liberation into the air. The physical shedding mechanism of airborne propagules in the 100--250 μm size range are not well understood. Many fungal plant pathogens have propagules in this size range that are shed from the bottom of infected leaves. If turbulent air flow is sufficient to liberate the sporocarps of fungi from leaves then the aerodynamic forces exerted must be sufficient to overcome adhesive forces. In this study I have sought to quantify the magnitude and direction of these aerodynamic forces and their causal flow fields with dynamically scaled physical models. I chose a genus of powdery mildew because maturation of the sporocarp entails morphological changes that lever the sporocarp further away from the leaf surface and out of the viscous boundary layer. Consequently I varied the sporocarp morphology, the boundary layer thickness, and the flow velocity as forces on models were measured with a transducer. Additionally I analyzed the fluid velocity around the models using PIV, which allowed for quantification of the relative importance of shear forces and pressure-gradient forces. The results suggest that forces from steady and unsteady wind alike are insufficient to explain liberation.

Dynamic gas temperature measurement system. Volume 2: Operation and program manual

NASA Technical Reports Server (NTRS)

Purpura, P. T.

1983-01-01

The hot section technology (HOST) dynamic gas temperature measurement system computer program acquires data from two type B thermocouples of different diameters. The analysis method determines the in situ value of an aerodynamic parameter T, containing the heat transfer coefficient from the transfer function of the two thermocouples. This aerodynamic parameter is used to compute a fequency response spectrum and compensate the dynamic portion of the signal of the smaller thermocouple. The calculations for the aerodynamic parameter and the data compensation technique are discussed. Compensated data are presented in either the time or frequency domain, time domain data as dynamic temperature vs time, or frequency domain data.

Feasibility study for a numerical aerodynamic simulation facility. Volume 2: Hardware specifications/descriptions

NASA Technical Reports Server (NTRS)

Green, F. M.; Resnick, D. R.

1979-01-01

An FMP (Flow Model Processor) was designed for use in the Numerical Aerodynamic Simulation Facility (NASF). The NASF was developed to simulate fluid flow over three-dimensional bodies in wind tunnel environments and in free space. The facility is applicable to studying aerodynamic and aircraft body designs. The following general topics are discussed in this volume: (1) FMP functional computer specifications; (2) FMP instruction specification; (3) standard product system components; (4) loosely coupled network (LCN) specifications/description; and (5) three appendices: performance of trunk allocation contention elimination (trace) method, LCN channel protocol and proposed LCN unified second level protocol.

Real-Time Aerodynamic Parameter Estimation without Air Flow Angle Measurements

NASA Technical Reports Server (NTRS)

Morelli, Eugene A.

2010-01-01

A technique for estimating aerodynamic parameters in real time from flight data without air flow angle measurements is described and demonstrated. The method is applied to simulated F-16 data, and to flight data from a subscale jet transport aircraft. Modeling results obtained with the new approach using flight data without air flow angle measurements were compared to modeling results computed conventionally using flight data that included air flow angle measurements. Comparisons demonstrated that the new technique can provide accurate aerodynamic modeling results without air flow angle measurements, which are often difficult and expensive to obtain. Implications for efficient flight testing and flight safety are discussed.

Micro air vehicle motion tracking and aerodynamic modeling

NASA Astrophysics Data System (ADS)

Uhlig, Daniel V.

Aerodynamic performance of small-scale fixed-wing flight is not well understood, and flight data are needed to gain a better understanding of the aerodynamics of micro air vehicles (MAVs) flying at Reynolds numbers between 10,000 and 30,000. Experimental studies have shown the aerodynamic effects of low Reynolds number flow on wings and airfoils, but the amount of work that has been conducted is not extensive and mostly limited to tests in wind and water tunnels. In addition to wind and water tunnel testing, flight characteristics of aircraft can be gathered through flight testing. The small size and low weight of MAVs prevent the use of conventional on-board instrumentation systems, but motion tracking systems that use off-board triangulation can capture flight trajectories (position and attitude) of MAVs with minimal onboard instrumentation. Because captured motion trajectories include minute noise that depends on the aircraft size, the trajectory results were verified in this work using repeatability tests. From the captured glide trajectories, the aerodynamic characteristics of five unpowered aircraft were determined. Test results for the five MAVs showed the forces and moments acting on the aircraft throughout the test flights. In addition, the airspeed, angle of attack, and sideslip angle were also determined from the trajectories. Results for low angles of attack (less than approximately 20 deg) showed the lift, drag, and moment coefficients during nominal gliding flight. For the lift curve, the results showed a linear curve until stall that was generally less than finite wing predictions. The drag curve was well described by a polar. The moment coefficients during the gliding flights were used to determine longitudinal and lateral stability derivatives. The neutral point, weather-vane stability and the dihedral effect showed some variation with different trim speeds (different angles of attack). In the gliding flights, the aerodynamic characteristics 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 envelope.

Overview of Sensitivity Analysis and Shape Optimization for Complex Aerodynamic Configurations

NASA Technical Reports Server (NTRS)

Newman, Perry A.; Newman, James C., III; Barnwell, Richard W.; Taylor, Arthur C., III; Hou, Gene J.-W.

1998-01-01

This paper presents a brief overview of some of the more recent advances in steady aerodynamic shape-design sensitivity analysis and optimization, based on advanced computational fluid dynamics. The focus here is on those methods particularly well- suited to the study of geometrically complex configurations and their potentially complex associated flow physics. When nonlinear state equations are considered in the optimization process, difficulties are found in the application of sensitivity analysis. Some techniques for circumventing such difficulties are currently being explored and are included here. Attention is directed to methods that utilize automatic differentiation to obtain aerodynamic sensitivity derivatives for both complex configurations and complex flow physics. Various examples of shape-design sensitivity analysis for unstructured-grid computational fluid dynamics algorithms are demonstrated for different formulations of the sensitivity equations. Finally, the use of advanced, unstructured-grid computational fluid dynamics in multidisciplinary analyses and multidisciplinary sensitivity analyses within future optimization processes is recommended and encouraged.

Aeroelastic coupling of geometrically nonlinear structures and linear unsteady aerodynamics: Two formulations

NASA Astrophysics Data System (ADS)

Demasi, L.; Livne, E.

2009-07-01

Two different time domain formulations of integrating commonly used frequency-domain unsteady aerodynamic models based on a modal approach with full order finite element models for structures with geometric nonlinearities are presented. Both approaches are tailored to flight vehicle configurations where geometric stiffness effects are important but where deformations are moderate, flow is attached, and linear unsteady aerodynamic modeling is adequate, such as low aspect ratio wings or joined-wing and strut-braced wings at small to moderate angles of attack. Results obtained using the two approaches are compared using both planar and non-planar wing configurations. Sub-critical and post-flutter speeds are considered. It is demonstrated that the two methods lead to the same steady solution for the sub-critical case after the transients subside. It is also shown that the two methods predict the amplitude and frequency of limit cycle oscillation (when present) with the same accuracy.

Aerodynamic performance of a small vertical axis wind turbine using an overset grid method

NASA Astrophysics Data System (ADS)

Bangga, Galih; Solichin, Mochammad; Daman, Aida; Sa'adiyah, Devy; Dessoky, Amgad; Lutz, Thorsten

2017-08-01

The present paper aims to asses the aerodynamic performance of a small vertical axis wind turbine operating at a small wind speed of 5 m/s for 6 different tip speed ratios (λ=2-7). The turbine consists of two blades constructed using the NACA 0015 airfoil. The study is carried out using computational fluid dynamics (CFD) methods employing an overset grid approach. The (URANS) SST k - ω is used as the turbulence model. For the preliminary study, simulations of the NACA 0015 under static conditions for a broad range of angle of attack and a rotating two-bladed VAWT are carried out. The results are compared with available measurement data and a good agreement is obtained. The simulations demonstrate that the maximum power coefficient attained is 0.45 for λ=4. The aerodynamic loads hysteresis are presented showing that the dynamic stall effect decreases with λ.

Environmental dynamics at orbital altitudes

NASA Technical Reports Server (NTRS)

Karr, G. R.

1976-01-01

The influence of real satellite aerodynamics on the determination of upper atmospheric density was investigated. A method of analysis of satellite drag data is presented which includes the effect of satellite lift and the variation in aerodynamic properties around the orbit. The studies indicate that satellite lift may be responsible for the observed orbit precession rather than a super rotation of the upper atmosphere. The influence of simplifying assumptions concerning the aerodynamics of objects in falling sphere analysis were evaluated and an improved method of analysis was developed. Wind tunnel data was used to develop more accurate drag coefficient relationships for studying altitudes between 80 and 120 Km. The improved drag coefficient relationships revealed a considerable error in previous falling sphere drag interpretation. These data were reanalyzed using the more accurate relationships. Theoretical investigations of the drag coefficient in the very low speed ratio region were also conducted.

Aerodynamic design and optimization in one shot

NASA Technical Reports Server (NTRS)

Ta'asan, Shlomo; Kuruvila, G.; Salas, M. D.

1992-01-01

This paper describes an efficient numerical approach for the design and optimization of aerodynamic bodies. As in classical optimal control methods, the present approach introduces a cost function and a costate variable (Lagrange multiplier) in order to achieve a minimum. High efficiency is achieved by using a multigrid technique to solve for all the unknowns simultaneously, but restricting work on a design variable only to grids on which their changes produce nonsmooth perturbations. Thus, the effort required to evaluate design variables that have nonlocal effects on the solution is confined to the coarse grids. However, if a variable has a nonsmooth local effect on the solution in some neighborhood, it is relaxed in that neighborhood on finer grids. The cost of solving the optimal control problem is shown to be approximately two to three times the cost of the equivalent analysis problem. Examples are presented to illustrate the application of the method to aerodynamic design and constraint optimization.

Integrated aerodynamic-structural design of a forward-swept transport wing

NASA Technical Reports Server (NTRS)

Haftka, Raphael T.; Grossman, Bernard; Kao, Pi-Jen; Polen, David M.; Sobieszczanski-Sobieski, Jaroslaw

1989-01-01

The introduction of composite materials is having a profound effect on aircraft design. Since these materials permit the designer to tailor material properties to improve structural, aerodynamic and acoustic performance, they require an integrated multidisciplinary design process. Futhermore, because of the complexity of the design process, numerical optimization methods are required. The utilization of integrated multidisciplinary design procedures for improving aircraft design is not currently feasible because of software coordination problems and the enormous computational burden. Even with the expected rapid growth of supercomputers and parallel architectures, these tasks will not be practical without the development of efficient methods for cross-disciplinary sensitivities and efficient optimization procedures. The present research is part of an on-going effort which is focused on the processes of simultaneous aerodynamic and structural wing design as a prototype for design integration. A sequence of integrated wing design procedures has been developed in order to investigate various aspects of the design process.

Aerodynamic Classification of Swept-Wing Ice Accretion

NASA Technical Reports Server (NTRS)

Diebold, Jeff M.; Broeren, Andy P.; Bragg, Michael B.

2013-01-01

The continued design, certification and safe operation of swept-wing airplanes in icing conditions rely on the advancement of computational and experimental simulation methods for higher fidelity results over an increasing range of aircraft configurations and performance, and icing conditions. The current stateof- the-art in icing aerodynamics is mainly built upon a comprehensive understanding of two-dimensional geometries that does not currently exist for fundamentally three-dimensional geometries such as swept wings. The purpose of this report is to describe what is known of iced-swept-wing aerodynamics and to identify the type of research that is required to improve the current understanding. Following the method used in a previous review of iced-airfoil aerodynamics, this report proposes a classification of swept-wing ice accretion into four groups based upon unique flowfield attributes. These four groups are: ice roughness, horn ice, streamwise ice and spanwise-ridge ice. In the case of horn ice it is shown that a further subclassification of "nominally 3D" or "highly 3D" horn ice may be necessary. For all of the proposed ice-shape classifications, relatively little is known about the three-dimensional flowfield and even less about the effect of Reynolds number and Mach number on these flowfields. The classifications and supporting data presented in this report can serve as a starting point as new research explores swept-wing aerodynamics with ice shapes. As further results are available, it is expected that these classifications will need to be updated and revised.

Aerodynamic Classification of Swept-Wing Ice Accretion

NASA Technical Reports Server (NTRS)

Diebold, Jeff M.; Broeren, Andy P.; Bragg, Michael B.

2013-01-01

The continued design, certification and safe operation of swept-wing airplanes in icing conditions rely on the advancement of computational and experimental simulation methods for higher fidelity results over an increasing range of aircraft configurations and performance, and icing conditions. The current state-of-the-art in icing aerodynamics is mainly built upon a comprehensive understanding of two-dimensional geometries that does not currently exist for fundamentally three-dimensional geometries such as swept wings. The purpose of this report is to describe what is known of iced-swept-wing aerodynamics and to identify the type of research that is required to improve the current understanding. Following the method used in a previous review of iced-airfoil aerodynamics, this report proposes a classification of swept-wing ice accretion into four groups based upon unique flowfield attributes. These four groups are: ice roughness, horn ice, streamwise ice and spanwise-ridge ice. In the case of horn ice it is shown that a further subclassification of nominally 3D or highly 3D horn ice may be necessary. For all of the proposed ice-shape classifications, relatively little is known about the three-dimensional flowfield and even less about the effect of Reynolds number and Mach number on these flowfields. The classifications and supporting data presented in this report can serve as a starting point as new research explores swept-wing aerodynamics with ice shapes. As further results are available, it is expected that these classifications will need to be updated and revised.

An influence coefficient method for the application of the modal technique to wing flutter suppression of the DAST ARW-1 wing

NASA Technical Reports Server (NTRS)

Pines, S.

1981-01-01

The methods used to compute the mass, structural stiffness, and aerodynamic forces in the form of influence coefficient matrices as applied to a flutter analysis of the Drones for Aerodynamic and Structural Testing (DAST) Aeroelastic Research Wing. The DAST wing was chosen because wind tunnel flutter test data and zero speed vibration data of the modes and frequencies exist and are available for comparison. A derivation of the equations of motion that can be used to apply the modal method for flutter suppression is included. A comparison of the open loop flutter predictions with both wind tunnel data and other analytical methods is presented.

Novel methodology for wide-ranged multistage morphing waverider based on conical theory

NASA Astrophysics Data System (ADS)

Liu, Zhen; Liu, Jun; Ding, Feng; Xia, Zhixun

2017-11-01

This study proposes the wide-ranged multistage morphing waverider design method. The flow field structure and aerodynamic characteristics of multistage waveriders are also analyzed. In this method, the multistage waverider is generated in the same conical flowfield, which contains a free-stream surface and different compression-stream surfaces. The obtained results show that the introduction of the multistage waverider design method can solve the problem of aerodynamic performance deterioration in the off-design state and allow the vehicle to always maintain the optimal flight state. The multistage waverider design method, combined with transfiguration flight strategy, can lead to greater design flexibility and the optimization of hypersonic wide-ranged waverider vehicles.

Summary of methods for calculating dynamic lateral stability and response and for estimating aerodynamic stability derivatives

NASA Technical Reports Server (NTRS)

Campbell, John P; Mckinney, Marion O

1952-01-01

A summary of methods for making dynamic lateral stability and response calculations and for estimating the aerodynamic stability derivatives required for use in these calculations is presented. The processes of performing calculations of the time histories of lateral motions, of the period and damping of these motions, and of the lateral stability boundaries are presented as a series of simple straightforward steps. Existing methods for estimating the stability derivatives are summarized and, in some cases, simple new empirical formulas are presented. Detailed estimation methods are presented for low-subsonic-speed conditions but only a brief discussion and a list of references are given for transonic and supersonic speed conditions.

Numerical simulation of aerodynamic characteristics of multi-element wing with variable flap

NASA Astrophysics Data System (ADS)

Lv, Hongyan; Zhang, Xinpeng; Kuang, Jianghong

2017-10-01

Based on the Reynolds averaged Navier-Stokes equation, the mesh generation technique and the geometric modeling method, the influence of the Spalart-Allmaras turbulence model on the aerodynamic characteristics is investigated. In order to study the typical configuration of aircraft, a similar DLR-F11 wing is selected. Firstly, the 3D model of wing is established, and the 3D model of plane flight, take-off and landing is established. The mesh structure of the flow field is constructed and the mesh is generated by mesh generation software. Secondly, by comparing the numerical simulation with the experimental data, the prediction of the aerodynamic characteristics of the multi section airfoil in takeoff and landing stage is validated. Finally, the two flap deflection angles of take-off and landing are calculated, which provide useful guidance for the aerodynamic characteristics of the wing and the flap angle design of the wing.

Method of Determining the Aerodynamic Characteristics of a Flying Vehicle from the Surface Pressure

NASA Astrophysics Data System (ADS)

Volkov, V. F.; Dyad'kin, A. A.; Zapryagaev, V. I.; Kiselev, N. P.

2017-11-01

The paper presents a description of the procedure used for determining the aerodynamic characteristics (forces and moments acting on a model of a flying vehicle) obtained from the results of pressure measurements on the surface of a model of a re-entry vehicle with operating retrofire brake rockets in the regime of hovering over a landing surface is given. The algorithm for constructing the interpolation polynomial over interpolation nodes in the radial and azimuthal directions using the assumption on the symmetry of pressure distribution over the surface is presented. The aerodynamic forces and moments at different tilts of the vehicle are obtained. It is shown that the aerodynamic force components acting on the vehicle in the regime of landing and caused by the action of the vertical velocity deceleration nozzle jets are negligibly small in comparison with the engine thrust.

Computer program for aerodynamic and blading design of multistage axial-flow compressors

NASA Technical Reports Server (NTRS)

Crouse, J. E.; Gorrell, W. T.

1981-01-01

A code for computing the aerodynamic design of a multistage axial-flow compressor and, if desired, the associated blading geometry input for internal flow analysis codes is presented. Compressible flow, which is assumed to be steady and axisymmetric, is the basis for a two-dimensional solution in the meridional plane with viscous effects modeled by pressure loss coefficients and boundary layer blockage. The radial equation of motion and the continuity equation are solved with the streamline curvature method on calculation stations outside the blade rows. The annulus profile, mass flow, pressure ratio, and rotative speed are input. A number of other input parameters specify and control the blade row aerodynamics and geometry. In particular, blade element centerlines and thicknesses can be specified with fourth degree polynomials for two segments. The output includes a detailed aerodynamic solution and, if desired, blading coordinates that can be used for internal flow analysis codes.

Aerodynamics of magnetic levitation (MAGLEV) trains

NASA Technical Reports Server (NTRS)

Schetz, Joseph A.; Marchman, James F., III

1996-01-01

High-speed (500 kph) trains using magnetic forces for levitation, propulsion and control offer many advantages for the nation and a good opportunity for the aerospace community to apply 'high tech' methods to the domestic sector. One area of many that will need advanced research is the aerodynamics of such MAGLEV (Magnetic Levitation) vehicles. There are important issues with regard to wind tunnel testing and the application of CFD to these devices. This talk will deal with the aerodynamic design of MAGLEV vehicles with emphasis on wind tunnel testing. The moving track facility designed and constructed in the 6 ft. Stability Wind Tunnel at Virginia Tech will be described. Test results for a variety of MAGLEV vehicle configurations will be presented. The last topic to be discussed is a Multi-disciplinary Design approach that is being applied to MAGLEV vehicle configuration design including aerodynamics, structures, manufacturability and life-cycle cost.

Computation of the stability derivatives via CFD and the sensitivity equations

NASA Astrophysics Data System (ADS)

Lei, Guo-Dong; Ren, Yu-Xin

2011-04-01

The method to calculate the aerodynamic stability derivates of aircrafts by using the sensitivity equations is extended to flows with shock waves in this paper. Using the newly developed second-order cell-centered finite volume scheme on the unstructured-grid, the unsteady Euler equations and sensitivity equations are solved simultaneously in a non-inertial frame of reference, so that the aerodynamic stability derivatives can be calculated for aircrafts with complex geometries. Based on the numerical results, behavior of the aerodynamic sensitivity parameters near the shock wave is discussed. Furthermore, the stability derivatives are analyzed for supersonic and hypersonic flows. The numerical results of the stability derivatives are found in good agreement with theoretical results for supersonic flows, and variations of the aerodynamic force and moment predicted by the stability derivatives are very close to those obtained by CFD simulation for both supersonic and hypersonic flows.

Preliminary aerodynamic design considerations for advanced laminar flow aircraft configurations

NASA Technical Reports Server (NTRS)

Johnson, Joseph L., Jr.; Yip, Long P.; Jordan, Frank L., Jr.

1986-01-01

Modern composite manufacturing methods have provided the opportunity for smooth surfaces that can sustain large regions of natural laminar flow (NLF) boundary layer behavior and have stimulated interest in developing advanced NLF airfoils and improved aircraft designs. Some of the preliminary results obtained in exploratory research investigations on advanced aircraft configurations at the NASA Langley Research Center are discussed. Results of the initial studies have shown that the aerodynamic effects of configuration variables such as canard/wing arrangements, airfoils, and pusher-type and tractor-type propeller installations can be particularly significant at high angles of attack. Flow field interactions between aircraft components were shown to produce undesirable aerodynamic effects on a wing behind a heavily loaded canard, and the use of properly designed wing leading-edge modifications, such as a leading-edge droop, offset the undesirable aerodynamic effects by delaying wing stall and providing increased stall/spin resistance with minimum degradation of laminar flow behavior.

New technology in turbine aerodynamics.

NASA Technical Reports Server (NTRS)

Glassman, A. J.; Moffitt, T. P.

1972-01-01

Cursory review of some recent work that has been done in turbine aerodynamic research. Topics discussed include the aerodynamic effect of turbine coolant, high work-factor (ratio of stage work to square of blade speed) turbines, and computer methods for turbine design and performance prediction. Experimental cooled-turbine aerodynamics programs using two-dimensional cascades, full annular cascades, and cold rotating turbine stage tests are discussed with some typical results presented. Analytically predicted results for cooled blade performance are compared to experimental results. The problems and some of the current programs associated with the use of very high work factors for fan-drive turbines of high-bypass-ratio engines are discussed. Computer programs have been developed for turbine design-point performance, off-design performance, supersonic blade profile design, and the calculation of channel velocities for subsonic and transonic flowfields. The use of these programs for the design and analysis of axial and radial turbines is discussed.

Methods to Determine the Deformation of the IRVE Hypersonic Inflatable Aerodynamic Decelerator

NASA Technical Reports Server (NTRS)

Young, William R.

2011-01-01

Small resonant targets used in conjunction with a microwave reflectometer to determine the deformation of the Hypersonic Inflatable Aerodynamic Decelerator (HIAD) during reentry are investigated. The reflectometer measures the distance to the targets and from this the HIAD deformation is determined. The HIAD is used by the Inflatable Reentry Vehicle Experiment (IRVE) which investigates the use of inflatable heat shields for atmospheric reentry. After several different microwave reflectometer systems were analyzed and compared it was determined that the most desirable for this application is the Frequency Doubling Target method.

Evaluation of Ares-I Control System Robustness to Uncertain Aerodynamics and Flex Dynamics

NASA Technical Reports Server (NTRS)

Jang, Jiann-Woei; VanTassel, Chris; Bedrossian, Nazareth; Hall, Charles; Spanos, Pol

2008-01-01

This paper discusses the application of robust control theory to evaluate robustness of the Ares-I control systems. Three techniques for estimating upper and lower bounds of uncertain parameters which yield stable closed-loop response are used here: (1) Monte Carlo analysis, (2) mu analysis, and (3) characteristic frequency response analysis. All three methods are used to evaluate stability envelopes of the Ares-I control systems with uncertain aerodynamics and flex dynamics. The results show that characteristic frequency response analysis is the most effective of these methods for assessing robustness.

Aerodynamic Design of a Propeller for High-Altitude Balloon Trajectory Control

NASA Technical Reports Server (NTRS)

Eppler, Richard; Somers, Dan M.

2012-01-01

The aerodynamic design of a propeller for the trajectory control of a high-altitude, scientific balloon has been performed using theoretical methods developed especially for such applications. The methods are described. Optimum, nonlinear chord and twist distributions have been developed in conjunction with the design of a family of airfoils, the SE403, SE404, and SE405, for the propeller. The very low Reynolds numbers along the propeller blade fall in a range that has yet to be rigorously investigated, either experimentally or theoretically.

Unsteady aerodynamic forces and torques on falling parallelograms in coupled tumbling-helical motions

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.

Unsteady aerodynamic forces and torques on falling parallelograms in coupled tumbling-helical motions.

PubMed

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.

Application of the generalized reduced gradient method to conceptual aircraft design

NASA Technical Reports Server (NTRS)

Gabriele, G. A.

1984-01-01

The complete aircraft design process can be broken into three phases of increasing depth: conceptual design, preliminary design, and detail design. Conceptual design consists primarily of developing general arrangements and selecting the configuration that optimally satisfies all mission requirements. The result of the conceptual phase is a conceptual baseline configuration that serves as the starting point for the preliminary design phase. The conceptual design of an aircraft involves a complex trade-off of many independent variables that must be investigated before deciding upon the basic configuration. Some of these variables are discrete (number of engines), some represent different configurations (canard vs conventional tail) and some may represent incorporation of new technologies (aluminum vs composite materials). At Lockheed-Georgia, the sizing program is known as GASP (Generalized Aircraft Sizing Program). GASP is a large program containing analysis modules covering the many different disciplines involved fin defining the aricraft, such as aerodynamics, structures, stability and control, mission performance, and cost. These analysis modules provide first-level estimates the aircraft properties that are derived from handbook, experimental, and historical sources.

MIST - MINIMUM-STATE METHOD FOR RATIONAL APPROXIMATION OF UNSTEADY AERODYNAMIC FORCE COEFFICIENT MATRICES

NASA Technical Reports Server (NTRS)

Karpel, M.

1994-01-01

Various control analysis, design, and simulation techniques of aeroservoelastic systems require the equations of motion to be cast in a linear, time-invariant state-space form. In order to account for unsteady aerodynamics, rational function approximations must be obtained to represent them in the first order equations of the state-space formulation. A computer program, MIST, has been developed which determines minimum-state approximations of the coefficient matrices of the unsteady aerodynamic forces. The Minimum-State Method facilitates the design of lower-order control systems, analysis of control system performance, and near real-time simulation of aeroservoelastic phenomena such as the outboard-wing acceleration response to gust velocity. Engineers using this program will be able to calculate minimum-state rational approximations of the generalized unsteady aerodynamic forces. Using the Minimum-State formulation of the state-space equations, they will be able to obtain state-space models with good open-loop characteristics while reducing the number of aerodynamic equations by an order of magnitude more than traditional approaches. These low-order state-space mathematical models are good for design and simulation of aeroservoelastic systems. The computer program, MIST, accepts tabular values of the generalized aerodynamic forces over a set of reduced frequencies. It then determines approximations to these tabular data in the LaPlace domain using rational functions. MIST provides the capability to select the denominator coefficients in the rational approximations, to selectably constrain the approximations without increasing the problem size, and to determine and emphasize critical frequency ranges in determining the approximations. MIST has been written to allow two types data weighting options. The first weighting is a traditional normalization of the aerodynamic data to the maximum unit value of each aerodynamic coefficient. The second allows weighting the importance of different tabular values in determining the approximations based upon physical characteristics of the system. Specifically, the physical weighting capability is such that each tabulated aerodynamic coefficient, at each reduced frequency value, is weighted according to the effect of an incremental error of this coefficient on aeroelastic characteristics of the system. In both cases, the resulting approximations yield a relatively low number of aerodynamic lag states in the subsequent state-space model. MIST is written in ANSI FORTRAN 77 for DEC VAX series computers running VMS. It requires approximately 1Mb of RAM for execution. The standard distribution medium for this package is a 9-track 1600 BPI magnetic tape in DEC VAX FILES-11 format. It is also available on a TK50 tape cartridge in DEC VAX BACKUP format. MIST was developed in 1991. DEC VAX and VMS are trademarks of Digital Equipment Corporation. FORTRAN 77 is a registered trademark of Lahey Computer Systems, Inc.

The equivalent angle-of-attack method for estimating the nonlinear aerodynamic characteristics of missile wings and control surfaces

NASA Technical Reports Server (NTRS)

Hemsch, M. J.; Nielsen, J. N.

1982-01-01

A method has been developed for estimating the nonlinear aerodynamic characteristics of missile wing and control surfaces. The method is based on the following assumption: if a fin on a body has the same normal-force coefficient as a wing alone composed of two of the same fins joined together at their root chords, then the other force and moment coefficients of the fin and the wing alone are the same including the nonlinearities. The method can be used for deflected fins at arbitrary bank angles and at high angles of attack. In the paper, a full derivation of the method is given, its accuracy demonstrated and its use in extending missile data bases is shown.

Survey of engineering computational methods and experimental programs for estimating supersonic missile aerodynamic characteristics

NASA Technical Reports Server (NTRS)

Sawyer, W. C.; Allen, J. M.; Hernandez, G.; Dillenius, M. F. E.; Hemsch, M. J.

1982-01-01

This paper presents a survey of engineering computational methods and experimental programs used for estimating the aerodynamic characteristics of missile configurations. Emphasis is placed on those methods which are suitable for preliminary design of conventional and advanced concepts. An analysis of the technical approaches of the various methods is made in order to assess their suitability to estimate longitudinal and/or lateral-directional characteristics for different classes of missile configurations. Some comparisons between the predicted characteristics and experimental data are presented. These comparisons are made for a large variation in flow conditions and model attitude parameters. The paper also presents known experimental research programs developed for the specific purpose of validating analytical methods and extending the capability of data-base programs.

Overview of the AVT-191 Project to Assess Sensitivity Analysis and Uncertainty Quantification Methods for Military Vehicle Design

NASA Technical Reports Server (NTRS)

Benek, John A.; Luckring, James M.

2017-01-01

A NATO symposium held in 2008 identified many promising sensitivity analysis and un-certainty quantification technologies, but the maturity and suitability of these methods for realistic applications was not known. The STO Task Group AVT-191 was established to evaluate the maturity and suitability of various sensitivity analysis and uncertainty quantification methods for application to realistic problems of interest to NATO. The program ran from 2011 to 2015, and the work was organized into four discipline-centric teams: external aerodynamics, internal aerodynamics, aeroelasticity, and hydrodynamics. This paper presents an overview of the AVT-191 program content.

Summary Findings from the AVT-191 Project to Assess Sensitivity Analysis and Uncertainty Quantification Methods for Military Vehicle Design

NASA Technical Reports Server (NTRS)

Benek, John A.; Luckring, James M.

2017-01-01

A NATO symposium held in Greece in 2008 identified many promising sensitivity analysis and uncertainty quantification technologies, but the maturity and suitability of these methods for realistic applications was not clear. The NATO Science and Technology Organization, Task Group AVT-191 was established to evaluate the maturity and suitability of various sensitivity analysis and uncertainty quantification methods for application to realistic vehicle development problems. The program ran from 2011 to 2015, and the work was organized into four discipline-centric teams: external aerodynamics, internal aerodynamics, aeroelasticity, and hydrodynamics. This paper summarizes findings and lessons learned from the task group.

Comparison of theoretically predicted lateral-directional aerodynamic characteristics with full-scale wind tunnel data on the ATLIT airplane

NASA Technical Reports Server (NTRS)

Griswold, M.; Roskam, J.

1980-01-01

An analytical method is presented for predicting lateral-directional aerodynamic characteristics of light twin engine propeller-driven airplanes. This method is applied to the Advanced Technology Light Twin Engine airplane. The calculated characteristics are correlated against full-scale wind tunnel data. The method predicts the sideslip derivatives fairly well, although angle of attack variations are not well predicted. Spoiler performance was predicted somewhat high but was still reasonable. The rudder derivatives were not well predicted, in particular the effect of angle of attack. The predicted dynamic derivatives could not be correlated due to lack of experimental data.

Supersonic turbulent boundary layers with periodic mechanical non-equilibrium

NASA Astrophysics Data System (ADS)

Ekoto, Isaac Wesley

Previous studies have shown that favorable pressure gradients reduce the turbulence levels and length scales in supersonic flow. Wall roughness has been shown to reduce the large-scales in wall bounded flow. Based on these previous observations new questions have been raised. The fundamental questions this dissertation addressed are: (1) What are the effects of wall topology with sharp versus blunt leading edges? and (2) Is it possible that a further reduction of turbulent scales can occur if surface roughness and favorable pressure gradients are combined? To answer these questions and to enhance the current experimental database, an experimental analysis was performed to provide high fidelity documentation of the mean and turbulent flow properties along with surface and flow visualizations of a high-speed (M = 2.86), high Reynolds number (Retheta ≈ 60,000) supersonic turbulent boundary layer distorted by curvature-induced favorable pressure gradients and large-scale ( k+s ≈ 300) uniform surface roughness. Nine models were tested at three separate locations. Three pressure gradient models strengths (a nominally zero, a weak, and a strong favorable pressure gradient) and three roughness topologies (aerodynamically smooth, square, and diamond shaped roughness elements) were used. Highly resolved planar measurements of mean and fluctuating velocity components were accomplished using particle image velocimetry. Stagnation pressure profiles were acquired with a traversing Pitot probe. Surface pressure distributions were characterized using pressure sensitive paint. Finally flow visualization was accomplished using schlieren photographs. Roughness topology had a significant effect on the boundary layer mean and turbulent properties due to shock boundary layer interactions. Favorable pressure gradients had the expected stabilizing effect on turbulent properties, but the improvements were less significant for models with surface roughness near the wall due to increased tendency towards flow separation. It was documented that proper roughness selection coupled with a sufficiently strong favorable pressure gradient produced regions of "negative" production in the transport of turbulent stress. This led to localized areas of significant turbulence stress reduction. With proper roughness selection and sufficient favorable pressure gradient strength, it is believed that localized relaminarization of the boundary layer is possible.

Dai-Kou type conjugate gradient methods with a line search only using gradient.

PubMed

Huang, Yuanyuan; Liu, Changhe

2017-01-01

In this paper, the Dai-Kou type conjugate gradient methods are developed to solve the optimality condition of an unconstrained optimization, they only utilize gradient information and have broader application scope. Under suitable conditions, the developed methods are globally convergent. Numerical tests and comparisons with the PRP+ conjugate gradient method only using gradient show that the methods are efficient.

Validation of 3-D Ice Accretion Measurement Methodology for Experimental Aerodynamic Simulation

NASA Technical Reports Server (NTRS)

Broeren, Andy P.; Addy, Harold E., Jr.; Lee, Sam; Monastero, Marianne C.

2015-01-01

Determining the adverse aerodynamic effects due to ice accretion often relies on dry-air wind-tunnel testing of artificial, or simulated, ice shapes. Recent developments in ice-accretion documentation methods have yielded a laser-scanning capability that can measure highly three-dimensional (3-D) features of ice accreted in icing wind tunnels. The objective of this paper was to evaluate the aerodynamic accuracy of ice-accretion simulations generated from laser-scan data. Ice-accretion tests were conducted in the NASA Icing Research Tunnel using an 18-in. chord, two-dimensional (2-D) straight wing with NACA 23012 airfoil section. For six ice-accretion cases, a 3-D laser scan was performed to document the ice geometry prior to the molding process. Aerodynamic performance testing was conducted at the University of Illinois low-speed wind tunnel at a Reynolds number of 1.8 × 10(exp 6) and a Mach number of 0.18 with an 18-in. chord NACA 23012 airfoil model that was designed to accommodate the artificial ice shapes. The ice-accretion molds were used to fabricate one set of artificial ice shapes from polyurethane castings. The laser-scan data were used to fabricate another set of artificial ice shapes using rapid prototype manufacturing such as stereolithography. The iced-airfoil results with both sets of artificial ice shapes were compared to evaluate the aerodynamic simulation accuracy of the laser-scan data. For five of the six ice-accretion cases, there was excellent agreement in the iced-airfoil aerodynamic performance between the casting and laser-scan based simulations. For example, typical differences in iced-airfoil maximum lift coefficient were less than 3 percent with corresponding differences in stall angle of approximately 1 deg or less. The aerodynamic simulation accuracy reported in this paper has demonstrated the combined accuracy of the laser-scan and rapid-prototype manufacturing approach to simulating ice accretion for a NACA 23012 airfoil. For several of the ice-accretion cases tested, the aerodynamics is known to depend upon the small, three-dimensional features of the ice. These data show that the laser-scan and rapid-prototype manufacturing approach is capable of replicating these ice features within the reported accuracies of the laser-scan measurement and rapid-prototyping method; thus providing a new capability for high-fidelity ice-accretion documentation and artificial ice-shape fabrication for icing research.

Validation of 3-D Ice Accretion Measurement Methodology for Experimental Aerodynamic Simulation

NASA Technical Reports Server (NTRS)

Broeren, Andy P.; Addy, Harold E., Jr.; Lee, Sam; Monastero, Marianne C.

2014-01-01

Determining the adverse aerodynamic effects due to ice accretion often relies on dry-air wind-tunnel testing of artificial, or simulated, ice shapes. Recent developments in ice accretion documentation methods have yielded a laser-scanning capability that can measure highly three-dimensional features of ice accreted in icing wind tunnels. The objective of this paper was to evaluate the aerodynamic accuracy of ice-accretion simulations generated from laser-scan data. Ice-accretion tests were conducted in the NASA Icing Research Tunnel using an 18-inch chord, 2-D straight wing with NACA 23012 airfoil section. For six ice accretion cases, a 3-D laser scan was performed to document the ice geometry prior to the molding process. Aerodynamic performance testing was conducted at the University of Illinois low-speed wind tunnel at a Reynolds number of 1.8 x 10(exp 6) and a Mach number of 0.18 with an 18-inch chord NACA 23012 airfoil model that was designed to accommodate the artificial ice shapes. The ice-accretion molds were used to fabricate one set of artificial ice shapes from polyurethane castings. The laser-scan data were used to fabricate another set of artificial ice shapes using rapid prototype manufacturing such as stereolithography. The iced-airfoil results with both sets of artificial ice shapes were compared to evaluate the aerodynamic simulation accuracy of the laser-scan data. For four of the six ice-accretion cases, there was excellent agreement in the iced-airfoil aerodynamic performance between the casting and laser-scan based simulations. For example, typical differences in iced-airfoil maximum lift coefficient were less than 3% with corresponding differences in stall angle of approximately one degree or less. The aerodynamic simulation accuracy reported in this paper has demonstrated the combined accuracy of the laser-scan and rapid-prototype manufacturing approach to simulating ice accretion for a NACA 23012 airfoil. For several of the ice-accretion cases tested, the aerodynamics is known to depend upon the small, three dimensional features of the ice. These data show that the laser-scan and rapid-prototype manufacturing approach is capable of replicating these ice features within the reported accuracies of the laser-scan measurement and rapid-prototyping method; thus providing a new capability for high-fidelity ice-accretion documentation and artificial ice-shape fabrication for icing research.

PAN AIR summary document (version 1.0)

NASA Technical Reports Server (NTRS)

Derbyshire, T.; Sidwell, K. W.

1982-01-01

The capabilities and limitations of the panel aerodynamics (PAN AIR) computer program system are summarized. This program uses a higher order panel method to solve boundary value problems involving the Prandtl-Glauert equation for subsonic and supersonic potential flows. Both aerodynamic and hydrodynamic problems can be solved using this modular software which is written for the CDC 6600 and 7600, and the CYBER 170 series computers.

Aerodynamic Validation of Emerging Projectile and Missile Configurations

DTIC Science & Technology

2010-12-01

Inflation Layers at the Surface of the M549 Projectile....................................39 Figure 33. Probe Profile from Nose to Shock Front...behavior is critical for the design of new projectile shapes. The conventional approach to predict this aerodynamic behavior is through wind tunnel ...tool to study fluid flows and complements empirical methods and wind tunnel testing. In this study, the computer program ANSYS CFX was used to

A system for aerodynamic design and analysis of supersonic aircraft. Part 4: Test cases

NASA Technical Reports Server (NTRS)

Middleton, W. D.; Lundry, J. L.

1980-01-01

An integrated system of computer programs was developed for the design and analysis of supersonic configurations. The system uses linearized theory methods for the calculation of surface pressures and supersonic area rule concepts in combination with linearized theory for calculation of aerodynamic force coefficients. Interactive graphics are optional at the user's request. Representative test cases and associated program output are presented.

Aerodynamics: The Wright Way

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.

A fast and accurate method to predict 2D and 3D aerodynamic boundary layer flows

NASA Astrophysics Data System (ADS)

Bijleveld, H. A.; Veldman, A. E. P.

2014-12-01

A quasi-simultaneous interaction method is applied to predict 2D and 3D aerodynamic flows. This method is suitable for offshore wind turbine design software as it is a very accurate and computationally reasonably cheap method. This study shows the results for a NACA 0012 airfoil. The two applied solvers converge to the experimental values when the grid is refined. We also show that in separation the eigenvalues remain positive thus avoiding the Goldstein singularity at separation. In 3D we show a flow over a dent in which separation occurs. A rotating flat plat is used to show the applicability of the method for rotating flows. The shown capabilities of the method indicate that the quasi-simultaneous interaction method is suitable for design methods for offshore wind turbine blades.

High Energy Boundary Conditions for a Cartesian Mesh Euler Solver

NASA Technical Reports Server (NTRS)

Pandya, Shishir; Murman, Scott; Aftosmis, Michael

2003-01-01

Inlets and exhaust nozzles are common place in the world of flight. Yet, many aerodynamic simulation packages do not provide a method of modelling such high energy boundaries in the flow field. For the purposes of aerodynamic simulation, inlets and exhausts are often fared over and it is assumed that the flow differences resulting from this assumption are minimal. While this is an adequate assumption for the prediction of lift, the lack of a plume behind the aircraft creates an evacuated base region thus effecting both drag and pitching moment values. In addition, the flow in the base region is often mis-predicted resulting in incorrect base drag. In order to accurately predict these quantities, a method for specifying inlet and exhaust conditions needs to be available in aerodynamic simulation packages. A method for a first approximation of a plume without accounting for chemical reactions is added to the Cartesian mesh based aerodynamic simulation package CART3D. The method consists of 3 steps. In the first step, a components approach where each triangle is assigned a component number is used. Here, a method for marking the inlet or exhaust plane triangles as separate components is discussed. In step two, the flow solver is modified to accept a reference state for the components marked inlet or exhaust. In the third step, the flow solver uses these separated components and the reference state to compute the correct flow condition at that triangle. The present method is implemented in the CART3D package which consists of a set of tools for generating a Cartesian volume mesh from a set of component triangulations. The Euler equations are solved on the resulting unstructured Cartesian mesh. The present methods is implemented in this package and its usefulness is demonstrated with two validation cases. A generic missile body is also presented to show the usefulness of the method on a real world geometry.

In-Flight Aeroelastic Stability of the Thermal Protection System on the NASA HIAD, Part II: Nonlinear Theory and Extended Aerodynamics

NASA Technical Reports Server (NTRS)

Goldman, Benjamin D.; Dowell, Earl H.; Scott, Robert C.

2015-01-01

Conical shell theory and a supersonic potential flow aerodynamic theory are used to study the nonlinear pressure buckling and aeroelastic limit cycle behavior of the thermal protection system for NASA's Hypersonic Inflatable Aerodynamic Decelerator. The structural model of the thermal protection system consists of an orthotropic conical shell of the Donnell type, resting on several circumferential elastic supports. Classical Piston Theory is used initially for the aerodynamic pressure, but was found to be insufficient at low supersonic Mach numbers. Transform methods are applied to the convected wave equation for potential flow, and a time-dependent aerodynamic pressure correction factor is obtained. The Lagrangian of the shell system is formulated in terms of the generalized coordinates for all displacements and the Rayleigh-Ritz method is used to derive the governing differential-algebraic equations of motion. Aeroelastic limit cycle oscillations and buckling deformations are calculated in the time domain using a Runge-Kutta method in MATLAB. Three conical shell geometries were considered in the present analysis: a 3-meter diameter 70 deg. cone, a 3.7-meter 70 deg. cone, and a 6-meter diameter 70 deg. cone. The 6-meter configuration was loaded statically and the results were compared with an experimental load test of a 6-meter HIAD. Though agreement between theoretical and experimental strains was poor, the circumferential wrinkling phenomena observed during the experiments was captured by the theory and axial deformations were qualitatively similar in shape. With Piston Theory aerodynamics, the nonlinear flutter dynamic pressures of the 3-meter configuration were in agreement with the values calculated using linear theory, and the limit cycle amplitudes were generally on the order of the shell thickness. The effect of axial tension was studied for this configuration, and increasing tension was found to decrease the limit cycle amplitudes when the circumferential elastic supports were neglected, but resulted in more complex behavior when the supports were included. The nominal flutter dynamic pressure of the 3.7-meter configuration was significantly lower than that of the 3-meter, and it was found that two sets of natural modes coalesce to flutter modes near the same dynamic pressure. This resulted in a significant drop in the limit cycle frequencies at higher dynamic pressures, where the flutter mode with the lower frequency becomes more critical. Pre-buckling pressure loads and the aerodynamic pressure correction factor were studied for all geometries, and these effects resulted in significantly lower flutter boundaries compared with Piston Theory alone. The maximum dynamic pressure predicted by aerodynamic simulations of a proposed 3.7-meter HIAD vehicle was still lower than any of the calculated flutter dynamic pressures, suggesting that aeroelastic effects for this vehicle are of little concern.