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.

Impact of aerodynamic resistance formulations used in two-source modeling of energy exchange from the soil and vegetation using land surface temperature

USDA-ARS?s Scientific Manuscript database

Application of the Two-Source Energy Balance (TSEB) Model using land surface temperature (LST) requires aerodynamic resistance parameterizations for the flux exchange above the canopy layer, within the canopy air space and at the soil/substrate surface. There are a number of aerodynamic resistance f...

Review of Our National Heritage of Launch Vehicles Using Aerodynamic Surfaces and Current Use of These by Other Nations. Part II; Center Director's Discretionary Fund Project Numbe

NASA Technical Reports Server (NTRS)

Barret, C.

1996-01-01

Marshall Space Flight Center has a rich heritage of launch vehicles that have used aerodynamic surfaces for flight stability and for flight control. Recently, due to the aft center-of-gravity (cg) locations on launch vehicles currently being studied, the need has arisen for the vehicle control augmentation that can be provided by these flight controls. Aerodynamic flight control can also reduce engine gimbaling requirements, provide actuator failure protection, enhance crew safety, and increase vehicle reliability and payload capability. As a starting point for the novel design of aerodynamic flight control augmentors for a Saturn class, aft cg launch vehicle, this report undertakes a review of our national heritage of launch vehicles using aerodynamic surfaces, along with a survey of current use of aerodynamic surfaces on large launch vehicles of other nations. This report presents one facet of Center Director's Discretionary Fund Project 93-05 and has a previous and subsequent companion publication.

Workshop on Aircraft Surface Representation for Aerodynamic Computation

NASA Technical Reports Server (NTRS)

Gregory, T. J. (Editor); Ashbaugh, J. (Editor)

1980-01-01

Papers and discussions on surface representation and its integration with aerodynamics, computers, graphics, wind tunnel model fabrication, and flow field grid generation are presented. Surface definition is emphasized.

A FORTRAN program for calculating three dimensional, inviscid and rotational flows with shock waves in axial compressor blade rows: User's manual

NASA Technical Reports Server (NTRS)

Thompkins, W. T., Jr.

1982-01-01

A FORTRAN-IV computer program was developed for the calculation of the inviscid transonic/supersonic flow field in a fully three dimensional blade passage of an axial compressor rotor or stator. Rotors may have dampers (part span shrouds). MacCormack's explicit time marching method is used to solve the unsteady Euler equations on a finite difference mesh. This technique captures shocks and smears them over several grid points. Input quantities are blade row geometry, operating conditions and thermodynamic quanities. Output quantities are three velocity components, density and internal energy at each mesh point. Other flow quanities are calculated from these variables. A short graphics package is included with the code, and may be used to display the finite difference grid, blade geometry and static pressure contour plots on blade to blade calculation surfaces or blade suction and pressure surfaces. The flow in a low aspect ratio transonic compressor was analyzed and compared with high response total pressure probe measurements and gas fluorescence static density measurements made in the MIT blowdown wind tunnel. These comparisons show that the computed flow fields accurately model the measured shock wave locations and overall aerodynamic performance.

Numerical study of aerodynamic effects on road vehicles lifting surfaces

NASA Astrophysics Data System (ADS)

Cernat, Mihail Victor; Cernat Bobonea, Andreea

2017-01-01

The aerodynamic performance analysis of road vehicles depends on the study of engine intake and cooling flow, internal ventilation, tire cooling, and overall external flow as the motion of air around a moving vehicle affects all of its components in one form or another. Due to the complex geometry of these, the aerodynamic interaction between the various body components is significant, resulting in vortex flow and lifting surface shapes. The present study, however focuses on the effects of external aerodynamics only, and in particular on the flow over the lifting surfaces of a common compact car, designed especially for this study.

Aerodynamic computation of gliders

NASA Technical Reports Server (NTRS)

Schrenk, M

1922-01-01

In the following discussion, a knowledge of the theoretical principles of airplane construction is assumed, as presented in detail by Vogt and Lippisch. A few quantities will however be otherwise designated, in accordance with the Gottingen symbols.

Development of Improved Surface Integral Methods for Jet Aeroacoustic Predictions

NASA Technical Reports Server (NTRS)

Pilon, Anthony R.; Lyrintzis, Anastasios S.

1997-01-01

The accurate prediction of aerodynamically generated noise has become an important goal over the past decade. Aeroacoustics must now be an integral part of the aircraft design process. The direct calculation of aerodynamically generated noise with CFD-like algorithms is plausible. However, large computer time and memory requirements often make these predictions impractical. It is therefore necessary to separate the aeroacoustics problem into two parts, one in which aerodynamic sound sources are determined, and another in which the propagating sound is calculated. This idea is applied in acoustic analogy methods. However, in the acoustic analogy, the determination of far-field sound requires the solution of a volume integral. This volume integration again leads to impractical computer requirements. An alternative to the volume integrations can be found in the Kirchhoff method. In this method, Green's theorem for the linear wave equation is used to determine sound propagation based on quantities on a surface surrounding the source region. The change from volume to surface integrals represents a tremendous savings in the computer resources required for an accurate prediction. This work is concerned with the development of enhancements of the Kirchhoff method for use in a wide variety of aeroacoustics problems. This enhanced method, the modified Kirchhoff method, is shown to be a Green's function solution of Lighthill's equation. It is also shown rigorously to be identical to the methods of Ffowcs Williams and Hawkings. This allows for development of versatile computer codes which can easily alternate between the different Kirchhoff and Ffowcs Williams-Hawkings formulations, using the most appropriate method for the problem at hand. The modified Kirchhoff method is developed primarily for use in jet aeroacoustics predictions. Applications of the method are shown for two dimensional and three dimensional jet flows. Additionally, the enhancements are generalized so that they may be used in any aeroacoustics problem.

Impacts of differing aerodynamic resistance formulae on modeled energy exchange at the above-canopy/within-canopy/soil interface

USDA-ARS?s Scientific Manuscript database

Application of the Two-Source Energy Balance (TSEB) Model using land surface temperature (LST) requires aerodynamic resistance parameterizations for the flux exchange above the canopy layer, within the canopy air space and at the soil/substrate surface. There are a number of aerodynamic resistance f...

Effect of static shape deformation on aerodynamics and aerothermodynamics of hypersonic inflatable aerodynamic decelerator

NASA Astrophysics Data System (ADS)

Guo, Jinghui; Lin, Guiping; Bu, Xueqin; Fu, Shiming; Chao, Yanmeng

2017-07-01

The inflatable aerodynamic decelerator (IAD), which allows heavier and larger payloads and offers flexibility in landing site selection at higher altitudes, possesses potential superiority in next generation space transport system. However, due to the flexibilities of material and structure assembly, IAD inevitably experiences surface deformation during atmospheric entry, which in turn alters the flowfield around the vehicle and leads to the variations of aerodynamics and aerothermodynamics. In the current study, the effect of the static shape deformation on the hypersonic aerodynamics and aerothermodynamics of a stacked tori Hypersonic Inflatable Aerodynamic Decelerator (HIAD) is demonstrated and analyzed in detail by solving compressible Navier-Stokes equations with Menter's shear stress transport (SST) turbulence model. The deformed shape is obtained by structural modeling in the presence of maximum aerodynamic pressure during entry. The numerical results show that the undulating shape deformation makes significant difference to flow structure. In particular, the more curved outboard forebody surface results in local flow separations and reattachments in valleys, which consequently yields remarkable fluctuations of surface conditions with pressure rising in valleys yet dropping on crests while shear stress and heat flux falling in valleys yet rising on crests. Accordingly, compared with the initial (undeformed) shape, the corresponding differences of surface conditions get more striking outboard, with maximum augmentations of 379 pa, 2224 pa, and 19.0 W/cm2, i.e., 9.8%, 305.9%, and 101.6% for the pressure, shear stress and heat flux respectively. Moreover, it is found that, with the increase of angle of attack, the aerodynamic characters and surface heating vary and the aeroheating disparities are evident between the deformed and initial shape. For the deformable HIAD model investigated in this study, the more intense surface conditions and changed flight aerodynamics are revealed, which is critical for the selection of structure material and design of flight control system.

Sensitivity Analysis for Coupled Aero-structural Systems

NASA Technical Reports Server (NTRS)

Giunta, Anthony A.

1999-01-01

A novel method has been developed for calculating gradients of aerodynamic force and moment coefficients for an aeroelastic aircraft model. This method uses the Global Sensitivity Equations (GSE) to account for the aero-structural coupling, and a reduced-order modal analysis approach to condense the coupling bandwidth between the aerodynamic and structural models. Parallel computing is applied to reduce the computational expense of the numerous high fidelity aerodynamic analyses needed for the coupled aero-structural system. Good agreement is obtained between aerodynamic force and moment gradients computed with the GSE/modal analysis approach and the same quantities computed using brute-force, computationally expensive, finite difference approximations. A comparison between the computational expense of the GSE/modal analysis method and a pure finite difference approach is presented. These results show that the GSE/modal analysis approach is the more computationally efficient technique if sensitivity analysis is to be performed for two or more aircraft design parameters.

Active control of multi-element rotor blade airfoils

NASA Technical Reports Server (NTRS)

Torok, Michael S. (Inventor); Moffitt, Robert C. (Inventor); Bagai, Ashish (Inventor)

2005-01-01

A multi-element rotor blade includes an individually controllable main element and fixed aerodynamic surface in an aerodynamically efficient location relative to the main element. The main element is controlled to locate the fixed aerodynamic surface in a position to increase lift and/or reduce drag upon the main element at various azimuthal positions during rotation.

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.

Relationships between aerodynamic roughness and land use and land cover in Baltimore, Maryland

USGS Publications Warehouse

Nicholas, F.W.; Lewis, J.E.

1980-01-01

Urbanization changes the radiative, thermal, hydrologic, and aerodynamic properties of the Earth's surface. Knowledge of these surface characteristics, therefore, is essential to urban climate analysis. Aerodynamic or surface roughness of urban areas is not well documented, however, because of practical constraints in measuring the wind profile in the presence of large buildings. Using an empirical method designed by Lettau, and an analysis of variance of surface roughness values calculated for 324 samples averaging 0.8 hectare (ha) of land use and land cover sample in Baltimore, Md., a strong statistical relation was found between aerodynamic roughness and urban land use and land cover types. Assessment of three land use and land cover systems indicates that some of these types have significantly different surface roughness characteristics. The tests further indicate that statistically significant differences exist in estimated surface roughness values when categories (classes) from different land use and land cover classification systems are used as surrogates. A Level III extension of the U.S. Geological Survey Level II land use and land cover classification system provided the most reliable results. An evaluation of the physical association between the aerodynamic properties of land use and land cover and the surface climate by numerical simulation of the surface energy balance indicates that changes in surface roughness within the range of values typical of the Level III categories induce important changes in the surface climate.

An asymptotic unsteady lifting-line theory with energetics and optimum motion of thrust-producing lifting surfaces. Thesis

NASA Technical Reports Server (NTRS)

Ahmadi, A. R.

1981-01-01

A low frequency unsteady lifting-line theory is developed for a harmonically oscillating wing of large aspect ratio. The wing is assumed to be chordwise rigid but completely flexible in the span direction. The theory is developed by use of the method of matched asymptotic expansions which reduces the problem from a singular integral equation to quadrature. The wing displacements are prescribed and the pressure field, airloads, and unsteady induced downwash are obtained in closed form. The influence of reduced frequency, aspect ratio, planform shape, and mode of oscillation on wing aerodynamics is demonstrated through numerical examples. Compared with lifting-surface theory, computation time is reduced significantly. Using the present theory, the energetic quantities associated with the propulsive performance of a finite wing oscillating in combined pitch and heave are obtained in closed form. Numerical examples are presented for an elliptic wing.

Direct launch using the electric rail gun

NASA Technical Reports Server (NTRS)

Barber, J. P.

1983-01-01

The concept explored involves using a large single stage electric rail gun to achieve orbital velocities. Exit aerodynamics, launch package design and size, interior ballistics, system and component sizing and design concepts are treated. Technology development status and development requirements are identified and described. The expense of placing payloads in Earth orbit using conventional chemical rockets is considerable. Chemical rockets are very inefficient in converting chemical energy into payload kinetic energy. A rocket motor is relatively expensive and is usually expended on each launch. In addition specialized and expensive forms of fuel are required. Gun launching payloads directly to orbit from the Earth's surface is a possible alternative. Guns are much more energy efficient than rockets. The high capital cost of the gun installation can be recovered by reusing it over and over again. Finally, relatively inexpensive fuel and large quantities of energy are readily available to a fixed installation on the Earth's surface.

Aerodynamic analysis of an isolated vehicle wheel

NASA Astrophysics Data System (ADS)

Leśniewicz, P.; Kulak, M.; Karczewski, M.

2014-08-01

Increasing fuel prices force the manufacturers to look into all aspects of car aerodynamics including wheels, tyres and rims in order to minimize their drag. By diminishing the aerodynamic drag of vehicle the fuel consumption will decrease, while driving safety and comfort will improve. In order to properly illustrate the impact of a rotating wheel aerodynamics on the car body, precise analysis of an isolated wheel should be performed beforehand. In order to represent wheel rotation in contact with the ground, presented CFD simulations included Moving Wall boundary as well as Multiple Reference Frame should be performed. Sliding mesh approach is favoured but too costly at the moment. Global and local flow quantities obtained during simulations were compared to an experiment in order to assess the validity of the numerical model. Results of investigation illustrates dependency between type of simulation and coefficients (drag and lift). MRF approach proved to be a better solution giving result closer to experiment. Investigation of the model with contact area between the wheel and the ground helps to illustrate the impact of rotating wheel aerodynamics on the car body.

Computational Aerodynamic Simulations of a Spacecraft Cabin Ventilation Fan Design

NASA Technical Reports Server (NTRS)

Tweedt, Daniel L.

2010-01-01

Quieter working environments for astronauts are needed if future long-duration space exploration missions are to be safe and productive. Ventilation and payload cooling fans are known to be dominant sources of noise, with the International Space Station being a good case in point. To address this issue cost effectively, early attention to fan design, selection, and installation has been recommended, leading to an effort by NASA to examine the potential for small-fan noise reduction by improving fan aerodynamic design. As a preliminary part of that effort, the aerodynamics of a cabin ventilation fan designed by Hamilton Sundstrand has been simulated using computational fluid dynamics codes, and the computed solutions analyzed to quantify various aspects of the fan aerodynamics and performance. Four simulations were performed at the design rotational speed: two at the design flow rate and two at off-design flow rates. Following a brief discussion of the computational codes, various aerodynamic- and performance-related quantities derived from the computed flow fields are presented along with relevant flow field details. The results show that the computed fan performance is in generally good agreement with stated design goals.

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.

Computation of transonic flow past projectiles at angle of attack

NASA Technical Reports Server (NTRS)

Reklis, R. P.; Sturek, W. B.; Bailey, F. R.

1978-01-01

Aerodynamic properties of artillery shell such as normal force and pitching moment reach peak values in a narrow transonic Mach number range. In order to compute these quantities, numerical techniques have been developed to obtain solutions to the three-dimensional transonic small disturbance equation about slender bodies at angle of attack. The computation is based on a plane relaxation technique involving Fourier transforms to partially decouple the three-dimensional difference equations. Particular care is taken to assure accurate solutions near corners found in shell designs. Computed surface pressures are compared to experimental measurements for circular arc and cone cylinder bodies which have been selected as test cases. Computed pitching moments are compared to range measurements for a typical projectile shape.

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.

Some aerodynamic considerations related to wind tunnel model surface definition

NASA Technical Reports Server (NTRS)

Gloss, B. B.

1980-01-01

The aerodynamic considerations related to model surface definition are examined with particular emphasis in areas of fabrication tolerances, model surface finish, and orifice induced pressure errors. The effect of model surface roughness texture on skin friction is also discussed. It is shown that at a given Reynolds number, any roughness will produce no skin friction penalty.

Small scale noise and wind tunnel tests of upper surface blowing nozzle flap concepts. Volume 1. Aerodynamic test results

NASA Technical Reports Server (NTRS)

Renselaer, D. J.; Nishida, R. S.; Wilkin, C. A.

1975-01-01

The results and analyses of aerodynamic and acoustic studies conducted on the small scale noise and wind tunnel tests of upper surface blowing nozzle flap concepts are presented. Various types of nozzle flap concepts were tested. These are an upper surface blowing concept with a multiple slot arrangement with seven slots (seven slotted nozzle), an upper surface blowing type with a large nozzle exit at approximately mid-chord location in conjunction with a powered trailing edge flap with multiple slots (split flow or partially slotted nozzle). In addition, aerodynamic tests were continued on a similar multi-slotted nozzle flap, but with 14 slots. All three types of nozzle flap concepts tested appear to be about equal in overall aerodynamic performance but with the split flow nozzle somewhat better than the other two nozzle flaps in the landing approach mode. All nozzle flaps can be deflected to a large angle to increase drag without significant loss in lift. The nozzle flap concepts appear to be viable aerodynamic drag modulation devices for landing.

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.

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.

Miniature Trailing Edge Effector for Aerodynamic Control

NASA Technical Reports Server (NTRS)

Lee, Hak-Tae (Inventor); Bieniawski, Stefan R. (Inventor); Kroo, Ilan M. (Inventor)

2008-01-01

Improved miniature trailing edge effectors for aerodynamic control are provided. Three types of devices having aerodynamic housings integrated to the trailing edge of an aerodynamic shape are presented, which vary in details of how the control surface can move. A bucket type device has a control surface which is the back part of a C-shaped member having two arms connected by the back section. The C-shaped section is attached to a housing at the ends of the arms, and is rotatable about an axis parallel to the wing trailing edge to provide up, down and neutral states. A flip-up type device has a control surface which rotates about an axis parallel to the wing trailing edge to provide up, down, neutral and brake states. A rotating type device has a control surface which rotates about an axis parallel to the chord line to provide up, down and neutral states.

Numerical study on influence of single control surface on aero elastic behavior of forward-swept wing

NASA Astrophysics Data System (ADS)

Wang, Ning; Su, Xinbing; Ma, Binlin; Zhang, Xiaofei

2017-10-01

In order to study the influence of elastic forward-swept wing (FSW) with single control surface, the computational fluid dynamics/computational structural dynamics (CFD/CSD) loose coupling static aero elastic numerical calculation method was adopted for numerical simulation. The effects of the elastic FSW with leading- or trailing-edge control surface on aero elastic characteristics were calculated and analysed under the condition of high subsonic speed. The result shows that, the deflection of every single control surface could change the aero elastic characteristics of elastic FSW greatly. Compared with the baseline model, when leading-edge control surface deflected up, under the condition of small angles of attack, the aerodynamic characteristics was poor, but the bending and torsional deformation decreased. Under the condition of moderate angles of attack, the aerodynamic characteristics was improved, but bending and torsional deformation increased; When leading-edge control surface deflected down, the aerodynamic characteristics was improved, the bending and torsional deformation decreased/increased under the condition of small/moderate angles of attack. Compared with the baseline model, when trailing-edge control surface deflected down, the aerodynamic characteristics was improved. The bending and torsional deformation increased under the condition of small angles of attack. The bending deformation increased under the condition of small angles of attack, but torsional deformation decreases under the condition of moderate angles of attack. So, for the elastic FSW, the deflection of trailing-edge control surface play a more important role on the improvement of aerodynamic and elastic deformation characteristics.

Aerodynamic Indicial Functions and Their Use in Aeroelastic Formulation of Lifting Surfaces

NASA Technical Reports Server (NTRS)

Marzocca, Piergiovanni; Librescu, Liviu; Silva, Walter A.

2000-01-01

An investigation related to the use of linear indicial functions in the time and frequency domains, enabling one to derive the proper aerodynamic loads as to study the subcritical response and flutter of swept lifting surfaces, respectively, of the open/closed loop aeroelastic system is presented. The expressions of the lift and aerodynamic moment in the frequency domain are given in terms of the Theodorsen's function, while, in the time domain, these are obtained directly with the help of the Wagner's function. Closed form solutions of aerodynamic derivatives are obtained, graphical representations are supplied and conclusions and prospects for further developments are outlined.

Vehicle underbody fairing

DOEpatents

Ortega, Jason M.; Salari, Kambiz; McCallen, Rose

2010-11-09

A vehicle underbody fairing apparatus for reducing aerodynamic drag caused by a vehicle wheel assembly, by reducing the size of a recirculation zone formed under the vehicle body immediately downstream of the vehicle wheel assembly. The fairing body has a tapered aerodynamic surface that extends from a front end to a rear end of the fairing body with a substantially U-shaped cross-section that tapers in both height and width. Fasteners or other mounting devices secure the fairing body to an underside surface of the vehicle body, so that the front end is immediately downstream of the vehicle wheel assembly and a bottom section of the tapered aerodynamic surface rises towards the underside surface as it extends in a downstream direction.

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.

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.

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.

Application of CAD/CAE class systems to aerodynamic analysis of electric race cars

NASA Astrophysics Data System (ADS)

Grabowski, L.; Baier, A.; Buchacz, A.; Majzner, M.; Sobek, M.

2015-11-01

Aerodynamics is one of the most important factors which influence on every aspect of a design of a car and car driving parameters. The biggest influence aerodynamics has on design of a shape of a race car body, especially when the main objective of the race is the longest distance driven in period of time, which can not be achieved without low energy consumption and low drag of a car. Designing shape of the vehicle body that must generate the lowest possible drag force, without compromising the other parameters of the drive. In the article entitled „Application of CAD/CAE class systems to aerodynamic analysis of electric race cars” are being presented problems solved by computer analysis of cars aerodynamics and free form modelling. Analysis have been subjected to existing race car of a Silesian Greenpower Race Team. On a basis of results of analysis of existence of Kammback aerodynamic effect innovative car body were modeled. Afterwards aerodynamic analysis were performed to verify existence of aerodynamic effect for innovative shape and to recognize aerodynamics parameters of the shape. Analysis results in the values of coefficients and aerodynamic drag forces. The resulting drag forces Fx, drag coefficients Cx(Cd) and aerodynamic factors Cx*A allowed to compare all of the shapes to each other. Pressure distribution, air velocities and streams courses were useful in determining aerodynamic features of analyzed shape. For aerodynamic tests was used Ansys Fluent CFD software. In a paper the ways of surface modeling with usage of Realize Shape module and classic surface modeling were presented. For shapes modeling Siemens NX 9.0 software was used. Obtained results were used to estimation of existing shapes and to make appropriate conclusions.

Review and assessment of turbulence models for hypersonic flows

NASA Astrophysics Data System (ADS)

Roy, Christopher J.; Blottner, Frederick G.

2006-10-01

Accurate aerodynamic prediction is critical for the design and optimization of hypersonic vehicles. Turbulence modeling remains a major source of uncertainty in the computational prediction of aerodynamic forces and heating for these systems. The first goal of this article is to update the previous comprehensive review of hypersonic shock/turbulent boundary-layer interaction experiments published in 1991 by Settles and Dodson (Hypersonic shock/boundary-layer interaction database. NASA CR 177577, 1991). In their review, Settles and Dodson developed a methodology for assessing experiments appropriate for turbulence model validation and critically surveyed the existing hypersonic experiments. We limit the scope of our current effort by considering only two-dimensional (2D)/axisymmetric flows in the hypersonic flow regime where calorically perfect gas models are appropriate. We extend the prior database of recommended hypersonic experiments (on four 2D and two 3D shock-interaction geometries) by adding three new geometries. The first two geometries, the flat plate/cylinder and the sharp cone, are canonical, zero-pressure gradient flows which are amenable to theory-based correlations, and these correlations are discussed in detail. The third geometry added is the 2D shock impinging on a turbulent flat plate boundary layer. The current 2D hypersonic database for shock-interaction flows thus consists of nine experiments on five different geometries. The second goal of this study is to review and assess the validation usage of various turbulence models on the existing experimental database. Here we limit the scope to one- and two-equation turbulence models where integration to the wall is used (i.e., we omit studies involving wall functions). A methodology for validating turbulence models is given, followed by an extensive evaluation of the turbulence models on the current hypersonic experimental database. A total of 18 one- and two-equation turbulence models are reviewed, and results of turbulence model assessments for the six models that have been extensively applied to the hypersonic validation database are compiled and presented in graphical form. While some of the turbulence models do provide reasonable predictions for the surface pressure, the predictions for surface heat flux are generally poor, and often in error by a factor of four or more. In the vast majority of the turbulence model validation studies we review, the authors fail to adequately address the numerical accuracy of the simulations (i.e., discretization and iterative error) and the sensitivities of the model predictions to freestream turbulence quantities or near-wall y+ mesh spacing. We recommend new hypersonic experiments be conducted which (1) measure not only surface quantities but also mean and fluctuating quantities in the interaction region and (2) provide careful estimates of both random experimental uncertainties and correlated bias errors for the measured quantities and freestream conditions. For the turbulence models, we recommend that a wide-range of turbulence models (including newer models) be re-examined on the current hypersonic experimental database, including the more recent experiments. Any future turbulence model validation efforts should carefully assess the numerical accuracy and model sensitivities. In addition, model corrections (e.g., compressibility corrections) should be carefully examined for their effects on a standard, low-speed validation database. Finally, as new experiments or direct numerical simulation data become available with information on mean and fluctuating quantities, they should be used to improve the turbulence models and thus increase their predictive capability.

Aerodynamic Evidence Pertaining to the Entry of Tektites into the Earth's Atmosphere

NASA Technical Reports Server (NTRS)

Chapman, Dean R.; Larson, Howard K.; Anderson, Lewis A.

1962-01-01

Evidence is presented which shows that the Australian and Java tektites entered the earth's atmosphere and experienced ablation by severe aerodynamic heating in hypervelocity flight. The laboratory experiments on hypervelocity ablation have reproduced ring-wave flow ridges and coiled circumferential flanges like those found on certain of these tektites. Systematic striae distortions exhibited in a thin layer beneath the front surface of australites also are reproduced in the laboratory ablation experiments, and are shown to correspond to the calculated distortions for aerodynamic ablation of a glass. About 98 percent of Australian tektites represent aerodynamically stable configurations during the ablative portion of an entry trajectory. Certain meteorites exhibit surface features similar to those on tektites.

Continued Development and Application of Circulation Control Pneumatic Technology to Advanced Transport Aircraft

NASA Technical Reports Server (NTRS)

Englar, Robert J.

1998-01-01

Personnel of the Georgia Tech Research Institute (GTRI) Aerospace and Transportation Lab have completed a four-year grant program to develop and evaluate the pneumatic aerodynamic technology known as Circulation Control (CC) or Circulation Control Wing (CCW) for advanced transport aircraft. This pneumatic technology, which employs low-level blowing from tangential slots over round or near-round trailing edges of airfoils, greatly augments the circulation around a lifting or control surface and thus enhances the aerodynamic forces and moments generated by that surface. Two-dimensional force augmentations as high as 80 times the input blowing momentum coefficient have been recorded experimentally for these blown devices, thus providing returns of 8000% on the jet momentum expended. A further benefit is the absence of moving parts such as mechanical flaps, slats, spoilers, ailerons, elevators and rudders from these pneumatic surfaces, or the use of only very small, simple, blown aerodynamic surfaces on synergistic designs which integrate the lift, drag and control surfaces. The application of these devices to advanced aircraft can offer significant benefits in their performance, efficiency, simplicity, reliability, economic cost of operation, noise reduction, and safety of flight. To further develop and evaluate this potential, this research effort was conducted by GTRI under grant for the NASA Langley Research Center, Applied Aerodynamics Division, Subsonic Aerodynamics Branch, between June 14, 1993 and May 31, 1997.

Measured wavenumber: frequency spectrum associated with acoustic and aerodynamic wall pressure fluctuations.

PubMed

Arguillat, Blandine; Ricot, Denis; Bailly, Christophe; Robert, Gilles

2010-10-01

Direct measurements of the wavenumber-frequency spectrum of wall pressure fluctuations beneath a turbulent plane channel flow have been performed in an anechoic wind tunnel. A rotative array has been designed that allows the measurement of a complete map, 63×63 measuring points, of cross-power spectral densities over a large area. An original post-processing has been developed to separate the acoustic and the aerodynamic exciting loadings by transforming space-frequency data into wavenumber-frequency spectra. The acoustic part has also been estimated from a simple Corcos-like model including the contribution of a diffuse sound field. The measured acoustic contribution to the surface pressure fluctuations is 5% of the measured aerodynamic surface pressure fluctuations for a velocity and boundary layer thickness relevant for automotive interior noise applications. This shows that for aerodynamically induced car interior noise, both contributions to the surface pressure fluctuations on car windows have to be taken into account.

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.

General Theory of Aerodynamic Instability and the Mechanism of Flutter

NASA Technical Reports Server (NTRS)

Theodorsen, Theodore

1979-01-01

The aerodynamic forces on an oscillating airfoil or airfoil-aileron combination of three independent degrees of freedom were determined. The problem resolves itself into the solution of certain definite integrals, which were identified as Bessel functions of the first and second kind, and of zero and first order. The theory, based on potential flow and the Kutta condition, is fundamentally equivalent to the conventional wing section theory relating to the steady case. The air forces being known, the mechanism of aerodynamic instability was analyzed. An exact solution, involving potential flow and the adoption of the Kutta condition, was derived. The solution is of a simple form and is expressed by means of an auxiliary parameter k. The flutter velocity, treated as the unknown quantity, was determined as a function of a certain ratio of the frequencies in the separate degrees of freedom for any magnitudes and combinations of the airfoil-aileron parameters.

CFD Assessment of Aerodynamic Degradation of a Subsonic Transport Due to Airframe Damage

NASA Technical Reports Server (NTRS)

Frink, Neal T.; Pirzadeh, Shahyar Z.; Atkins, Harold L.; Viken, Sally A.; Morrison, Joseph H.

2010-01-01

A computational study is presented to assess the utility of two NASA unstructured Navier-Stokes flow solvers for capturing the degradation in static stability and aerodynamic performance of a NASA General Transport Model (GTM) due to airframe damage. The approach is to correlate computational results with a substantial subset of experimental data for the GTM undergoing progressive losses to the wing, vertical tail, and horizontal tail components. The ultimate goal is to advance the probability of inserting computational data into the creation of advanced flight simulation models of damaged subsonic aircraft in order to improve pilot training. Results presented in this paper demonstrate good correlations with slope-derived quantities, such as pitch static margin and static directional stability, and incremental rolling moment due to wing damage. This study further demonstrates that high fidelity Navier-Stokes flow solvers could augment flight simulation models with additional aerodynamic data for various airframe damage scenarios.

Aerodynamic Analysis of Simulated Heat Shield Recession for the Orion Command Module

NASA Technical Reports Server (NTRS)

Bibb, Karen L.; Alter, Stephen J.; Mcdaniel, Ryan D.

2008-01-01

The aerodynamic effects of the recession of the ablative thermal protection system for the Orion Command Module of the Crew Exploration Vehicle are important for the vehicle guidance. At the present time, the aerodynamic effects of recession being handled within the Orion aerodynamic database indirectly with an additional safety factor placed on the uncertainty bounds. This study is an initial attempt to quantify the effects for a particular set of recessed geometry shapes, in order to provide more rigorous analysis for managing recession effects within the aerodynamic database. The aerodynamic forces and moments for the baseline and recessed geometries were computed at several trajectory points using multiple CFD codes, both viscous and inviscid. The resulting aerodynamics for the baseline and recessed geometries were compared. The forces (lift, drag) show negligible differences between baseline and recessed geometries. Generally, the moments show a difference between baseline and recessed geometries that correlates with the maximum amount of recession of the geometry. The difference between the pitching moments for the baseline and recessed geometries increases as Mach number decreases (and the recession is greater), and reach a value of -0.0026 for the lowest Mach number. The change in trim angle of attack increases from approx. 0.5deg at M = 28.7 to approx. 1.3deg at M = 6, and is consistent with a previous analysis with a lower fidelity engineering tool. This correlation of the present results with the engineering tool results supports the continued use of the engineering tool for future work. The present analysis suggests there does not need to be an uncertainty due to recession in the Orion aerodynamic database for the force quantities. The magnitude of the change in pitching moment due to recession is large enough to warrant inclusion in the aerodynamic database. An increment in the uncertainty for pitching moment could be calculated from these results and included in the development of the aerodynamic database uncertainty for pitching moment.

Aerodynamic Effects and Modeling of Damage to Transport Aircraft

NASA Technical Reports Server (NTRS)

Shah, Gautam H.

2008-01-01

A wind tunnel investigation was conducted to measure the aerodynamic effects of damage to lifting and stability/control surfaces of a commercial transport aircraft configuration. The modeling of such effects is necessary for the development of flight control systems to recover aircraft from adverse, damage-related loss-of-control events, as well as for the estimation of aerodynamic characteristics from flight data under such conditions. Damage in the form of partial or total loss of area was applied to the wing, horizontal tail, and vertical tail. Aerodynamic stability and control implications of damage to each surface are presented, to aid in the identification of potential boundaries in recoverable stability or control degradation. The aerodynamic modeling issues raised by the wind tunnel results are discussed, particularly the additional modeling requirements necessitated by asymmetries due to damage, and the potential benefits of such expanded modeling.

Real-Time Unsteady Loads Measurements Using Hot-Film Sensors

NASA Technical Reports Server (NTRS)

Mangalam, Arun S.; Moes, Timothy R.

2004-01-01

Several flight-critical aerodynamic problems such as buffet, flutter, stall, and wing rock are strongly affected or caused by abrupt changes in unsteady aerodynamic loads and moments. Advanced sensing and flow diagnostic techniques have made possible simultaneous identification and tracking, in realtime, of the critical surface, viscosity-related aerodynamic phenomena under both steady and unsteady flight conditions. The wind tunnel study reported here correlates surface hot-film measurements of leading edge stagnation point and separation point, with unsteady aerodynamic loads on a NACA 0015 airfoil. Lift predicted from the correlation model matches lift obtained from pressure sensors for an airfoil undergoing harmonic pitchup and pitchdown motions. An analytical model was developed that demonstrates expected stall trends for pitchup and pitchdown motions. This report demonstrates an ability to obtain unsteady aerodynamic loads in real time, which could lead to advances in air vehicle safety, performance, ride-quality, control, and health management.

Real-Time Unsteady Loads Measurements Using Hot-Film Sensors

NASA Technical Reports Server (NTRS)

Mangalam, Arun S.; Moes, Timothy R.

2004-01-01

Several flight-critical aerodynamic problems such as buffet, flutter, stall, and wing rock are strongly affected or caused by abrupt changes in unsteady aerodynamic loads and moments. Advanced sensing and flow diagnostic techniques have made possible simultaneous identification and tracking, in real-time, of the critical surface, viscosity-related aerodynamic phenomena under both steady and unsteady flight conditions. The wind tunnel study reported here correlates surface hot-film measurements of leading edge stagnation point and separation point, with unsteady aerodynamic loads on a NACA 0015 airfoil. Lift predicted from the correlation model matches lift obtained from pressure sensors for an airfoil undergoing harmonic pitchup and pitchdown motions. An analytical model was developed that demonstrates expected stall trends for pitchup and pitchdown motions. This report demonstrates an ability to obtain unsteady aerodynamic loads in real-time, which could lead to advances in air vehicle safety, performance, ride-quality, control, and health management.

Aerodynamic Performance Predictions of Single and Twin Jet Afterbodies

NASA Technical Reports Server (NTRS)

Carlson, John R.; Pao, S. Paul; Abdol-Hamid, Khaled S.; Jones, William T.

1995-01-01

The multiblock three-dimensional Navier-Stokes method PAB3D was utilized by the Component Integration Branch (formerly Propulsion Aerodynamics Branch) at the NASA-Langley Research Center in an international study sponsored by AGARD Working Group #17 for the assessment of the state-of-the-art of propulsion-airframe integration testing techniques and CFD prediction technologies. Three test geometries from ONERA involving fundamental flow physics and four geometries from NASA-LaRC involving realistic flow interactions of wing, body, tail, and jet plumes were chosen by the Working Group. An overview of results on four (1 ONERA and 3 LaRC) of the seven test cases is presented. External static pressures, integrated pressure drag and total drag were calculated for the Langley test cases and jet plume velocity profiles and turbulent viscous stresses were calculated for the ONERA test case. Only selected data from these calculations are presented in this paper. The complete data sets calculated by the participants will be presented in an AGARD summary report. Predicted surface static pressures compared favorably with experimental data for the Langley geometries. Predicted afterbody drag compared well with experiment. Predicted nozzle drag was typically low due to over-compression of the flow near the trailing edge. Total drag was typically high. Predicted jet plume quantities on the ONERA case compared generally well with data.

Aerodynamic Flow Control by Thermoacoustic Excitation from the Constituent Nanomaterials on the Platform Surface

DTIC Science & Technology

2016-02-01

Nanomaterials on the Platform Surface by Bryan Glaz Approved for public release; distribution is unlimited...Research Laboratory Aerodynamic Flow Control by Thermoacoustic Excitation from the Constituent Nanomaterials on the Platform Surface by Bryan Glaz...shall be subject to any penalty for failing to comply with a collection of information if it does not display a currently valid OMB control number

The compressible aerodynamics of rotating blades based on an acoustic formulation

NASA Technical Reports Server (NTRS)

Long, L. N.

1983-01-01

An acoustic formula derived for the calculation of the noise of moving bodies is applied to aerodynamic problems. The acoustic formulation is a time domain result suitable for slender wings and bodies moving at subsonic speeds. A singular integral equation is derived in terms of the surface pressure which must then be solved numerically for aerodynamic purposes. However, as the 'observer' is moved onto the body surface, the divergent integrals in the acoustic formulation are semiconvergent. The procedure for regularization (or taking principal values of divergent integrals) is explained, and some numerical examples for ellipsoids, wings, and lifting rotors are presented. The numerical results show good agreement with available measured surface pressure data.

Sensitivity Analysis and Optimization of Aerodynamic Configurations with Blend Surfaces

NASA Technical Reports Server (NTRS)

Thomas, A. M.; Tiwari, S. N.

1997-01-01

A novel (geometrical) parametrization procedure using solutions to a suitably chosen fourth order partial differential equation is used to define a class of airplane configurations. Inclusive in this definition are surface grids, volume grids, and grid sensitivity. The general airplane configuration has wing, fuselage, vertical tail and horizontal tail. The design variables are incorporated into the boundary conditions, and the solution is expressed as a Fourier series. The fuselage has circular cross section, and the radius is an algebraic function of four design parameters and an independent computational variable. Volume grids are obtained through an application of the Control Point Form method. A graphic interface software is developed which dynamically changes the surface of the airplane configuration with the change in input design variable. The software is made user friendly and is targeted towards the initial conceptual development of any aerodynamic configurations. Grid sensitivity with respect to surface design parameters and aerodynamic sensitivity coefficients based on potential flow is obtained using an Automatic Differentiation precompiler software tool ADIFOR. Aerodynamic shape optimization of the complete aircraft with twenty four design variables is performed. Unstructured and structured volume grids and Euler solutions are obtained with standard software to demonstrate the feasibility of the new surface definition.

Aerodynamics of a Party Balloon

ERIC Educational Resources Information Center

Cross, Rod

2007-01-01

It is well-known that a party balloon can be made to fly erratically across a room, but it can also be used for quantitative measurements of other aspects of aerodynamics. Since a balloon is light and has a large surface area, even relatively weak aerodynamic forces can be readily demonstrated or measured in the classroom. Accurate measurements…

Supersonic Aerodynamic Characteristics of Proposed Mars '07 Smart Lander Configurations

NASA Technical Reports Server (NTRS)

Murphy, Kelly J.; Horvath, Thomas J.; Erickson, Gary E.; Green, Joseph M.

2002-01-01

Supersonic aerodynamic data were obtained for proposed Mars '07 Smart Lander configurations in NASA Langley Research Center's Unitary Plan Wind Tunnel. The primary objective of this test program was to assess the supersonic aerodynamic characteristics of the baseline Smart Lander configuration with and without fixed shelf/tab control surfaces. Data were obtained over a Mach number range of 2.3 to 4.5, at a free stream Reynolds Number of 1 x 10(exp 6) based on body diameter. All configurations were run at angles of attack from -5 to 20 degrees and angles of sideslip of -5 to 5 degrees. These results were complemented with computational fluid dynamic (CFD) predictions to enhance the understanding of experimentally observed aerodynamic trends. Inviscid and viscous full model CFD solutions compared well with experimental results for the baseline and 3 shelf/tab configurations. Over the range tested, Mach number effects were shown to be small on vehicle aerodynamic characteristics. Based on the results from 3 different shelf/tab configurations, a fixed control surface appears to be a feasible concept for meeting aerodynamic performance metrics necessary to satisfy mission requirements.

Commercial turbofan engine exhaust nozzle flow analyses using PAB3D

NASA Technical Reports Server (NTRS)

Abdol-Hamid, Khaled S.; Uenishi, K.; Carlson, John R.; Keith, B. D.

1992-01-01

Recent developments of a three-dimensional (PAB3D) code have paved the way for a computational investigation of complex aircraft aerodynamic components. The PAB3D code was developed for solving the simplified Reynolds Averaged Navier-Stokes equations in a three-dimensional multiblock/multizone structured mesh domain. The present analysis was applied to commercial turbofan exhaust flow systems. Solution sensitivity to grid density is presented. Laminar flow solutions were developed for all grids and two-equation k-epsilon solutions were developed for selected grids. Static pressure distributions, mass flow and thrust quantities were calculated for on-design engine operating conditions. Good agreement between predicted surface static pressures and experimental data was observed at different locations. Mass flow was predicted within 0.2 percent of experimental data. Thrust forces were typically within 0.4 percent of experimental data.

Biologically Inspired mm-size Gliding UAV

NASA Astrophysics Data System (ADS)

Weihs, Daniel; Zussman, Eyal; Yarin, Alexander

2002-11-01

We present a first design of an unmanned aerial vehicle whose aerodynamic loads are carried by comb-like permeable surfaces. This concept was based on observation and analysis of the flight capabilities of the Thrip family of insects, whose have wings of this form and various plant seeds which use this concept to form an aerodynamic decelerator. This concept is only practical for at low Reynolds numbers, as the viscous trace of bodies moving in fluid becomes thicker as Re becomes smaller. When Re<1, a cylinder moving in fluid can drag along fluid up to several cylinder diameters. This drag effect can be used to produce a comb structure, where the ratio of comb rod to inter-rod distance can be up to 0.1 without significant throughflow, i.e. the comb is equivalent to a continuous surface. To demonstrate this principle, we have built artificial ``dandelion seeds", i.e. stable aerodynamic decelerators (parachutes) made of permeable light mats of submicron diameter nanofibers. We produced the nanofiber matrices (mats) by electrospinning of polymer solutions, obtaining fibers of 200-400 nm diameter. These fibers were then deposited on frames that serve to define the aerodynamic surfaces, thus producing the mat, with controllable density. For stability, the aerodynamic surfaces had positive dihedral ( inverted umbrella) forms with the fuselage hanging below. When dropped, the platforms fell freely through the air, apex down, reaching terminal velocity very quickly. By comparing the sink rate of the permeable structures with equivalent decelerators with continuous (Saran-wrap) surfaces we show that the permeable surfaces are equivalent to continuous surfaces, with significant weight savings, as long as the local Reynolds number is o(1).

Estimation of Supersonic Stage Separation Aerodynamics of Winged-Body Launch Vehicles Using Response Surface Methods

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

2010-01-01

Response surface methodology was used to estimate the longitudinal stage separation aerodynamic characteristics of a generic, bimese, winged multi-stage launch vehicle configuration at supersonic speeds in the NASA LaRC Unitary Plan Wind Tunnel. The Mach 3 staging was dominated by shock wave interactions between the orbiter and booster vehicles throughout the relative spatial locations of interest. The inference space was partitioned into several contiguous regions within which the separation aerodynamics were presumed to be well-behaved and estimable using central composite designs capable of fitting full second-order response functions. The underlying aerodynamic response surfaces of the booster vehicle in belly-to-belly proximity to the orbiter vehicle were estimated using piecewise-continuous lower-order polynomial functions. The quality of fit and prediction capabilities of the empirical models were assessed in detail, and the issue of subspace boundary discontinuities was addressed. Augmenting the central composite designs to full third-order using computer-generated D-optimality criteria was evaluated. The usefulness of central composite designs, the subspace sizing, and the practicality of fitting lower-order response functions over a partitioned inference space dominated by highly nonlinear and possibly discontinuous shock-induced aerodynamics are discussed.

Aerodynamic characteristics of two 10-percent-thick NASA supercritical airfoils with different upper surface curvature distributions. [Langley 8 foot transonic tunnel tests

NASA Technical Reports Server (NTRS)

Harris, C. D.

1974-01-01

In order to assess the degree to which the characteristic region of low curvature of the supercritical airfoil can be practically extended on the upper surface, the aerodynamic characteristics of two supercritical airfoils with different upper surface curvature distributions were measured at Mach numbers from 0.60 to 0.81. Integrated section force and moment data, surface pressure distributions, and typical wake survey profiles are presented.

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.

Aerodynamic shape optimization of a HSCT type configuration with improved surface definition

NASA Technical Reports Server (NTRS)

Thomas, Almuttil M.; Tiwari, Surendra N.

1994-01-01

Two distinct parametrization procedures of generating free-form surfaces to represent aerospace vehicles are presented. The first procedure is the representation using spline functions such as nonuniform rational b-splines (NURBS) and the second is a novel (geometrical) parametrization using solutions to a suitably chosen partial differential equation. The main idea is to develop a surface which is more versatile and can be used in an optimization process. Unstructured volume grid is generated by an advancing front algorithm and solutions obtained using an Euler solver. Grid sensitivity with respect to surface design parameters and aerodynamic sensitivity coefficients based on potential flow is obtained using an automatic differentiator precompiler software tool. Aerodynamic shape optimization of a complete aircraft with twenty four design variables is performed. High speed civil transport aircraft (HSCT) configurations are targeted to demonstrate the process.

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.

Heat transfer characteristics of hypersonic waveriders with an emphasis on leading edge effects

NASA Technical Reports Server (NTRS)

Vanmol, Denis O.; Anderson, John D., Jr.

1992-01-01

The present analysis of the heat-transfer characteristics of a family of viscous-optimized, 60 m-long waverider hypersonic vehicles gives attention to the transition from laminar to turbulent flow, and to how the transition affects aerodynamic heating distributions over the waverider surface. Two different constant-dynamic-pressure flight trajectories are considered, at 0.2 and 1.0 freestream atmospheres. For Mach numbers below 10, it is found that passive radiative cooling of the surface is sufficient. The degree of leading-edge bluntness required by aerodynamic heating constraints does not significantly degrade the aerodynamic performance of these waveriders.

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.

Effect of gas injection on drag and surface heat transfer rates for a 30° semi-apex angle blunt body flying at Mach 5.75

NASA Astrophysics Data System (ADS)

Sahoo, N.; Kulkarni, V.; Jagadeesh, G.; Reddy, K. P. J.

Effect of coolant gas injection in the stagnation region on the surface heat transfer rates and aerodynamic drag for a large angle blunt body flying at hypersonic Mach number is reported for two stagnation enthalpies. A 60° apex-angle blunt cone model is employed for this purpose with air injection at the nose through a hole of 2mm diameter. The convective surface heating rates and aerodynamic drag are measured simultaneously using surface mounted platinum thin film sensors and internally mounted accelerometer balance system, respectively. About 35-40% reduction in surface heating rates is observed in the vicinity of stagnation region whereas 15-25% reduction in surface heating rates is felt beyond the stagnation region at stagnation enthalpy of 1.6MJ/kg. The aerodynamic drag expressed in terms of drag coefficient is found to increase by 20% due to the air injection.

On simple aerodynamic sensitivity derivatives for use in interdisciplinary optimization

NASA Technical Reports Server (NTRS)

Doggett, Robert V., Jr.

1991-01-01

Low-aspect-ratio and piston aerodynamic theories are reviewed as to their use in developing aerodynamic sensitivity derivatives for use in multidisciplinary optimization applications. The basic equations relating surface pressure (or lift and moment) to normal wash are given and discussed briefly for each theory. The general means for determining selected sensitivity derivatives are pointed out. In addition, some suggestions in very general terms are included as to sample problems for use in studying the process of using aerodynamic sensitivity derivatives in optimization studies.

An analysis of aerodynamic requirements for coordinated bank-to-turn autopilots

NASA Technical Reports Server (NTRS)

Arrow, A.

1982-01-01

Two planar missile airframes were compared having the potential for improved bank-to-turn control but having different aerodynamic properties. The comparison was made with advanced level autopilots using both linear and nonlinear 3-D aerodynamic models to obtain realistic missile body angular rates and control surface incidence. Cortical cross-coupling effects are identified and desirable aerodynamics are recommended for improved coordinated (BTT) (CBTT) performance. In addition, recommendations are made for autopilot control law analyses and design techniques for improving CBTT performance.

Distributed Aerodynamic Sensing and Processing Toolbox

NASA Technical Reports Server (NTRS)

Brenner, Martin; Jutte, Christine; Mangalam, Arun

2011-01-01

A Distributed Aerodynamic Sensing and Processing (DASP) toolbox was designed and fabricated for flight test applications with an Aerostructures Test Wing (ATW) mounted under the fuselage of an F-15B on the Flight Test Fixture (FTF). DASP monitors and processes the aerodynamics with the structural dynamics using nonintrusive, surface-mounted, hot-film sensing. This aerodynamic measurement tool benefits programs devoted to static/dynamic load alleviation, body freedom flutter suppression, buffet control, improvement of aerodynamic efficiency through cruise control, supersonic wave drag reduction through shock control, etc. This DASP toolbox measures local and global unsteady aerodynamic load distribution with distributed sensing. It determines correlation between aerodynamic observables (aero forces) and structural dynamics, and allows control authority increase through aeroelastic shaping and active flow control. It offers improvements in flutter suppression and, in particular, body freedom flutter suppression, as well as aerodynamic performance of wings for increased range/endurance of manned/ unmanned flight vehicles. Other improvements include inlet performance with closed-loop active flow control, and development and validation of advanced analytical and computational tools for unsteady aerodynamics.

Aerodynamic characteristics of a canard-controlled missile at Mach numbers of 0.8, 1.3, and 1.75. [in the Ames 6 by 6 foot wind tunnel

NASA Technical Reports Server (NTRS)

Kassner, D. L.; Wettlaufer, B.

1977-01-01

A typical missile model with nose-mounted canards and cruciform tail surfaces was tested in the Ames 6- by 6-Foot Wind Tunnel to determine the contributions of the component aerodynamic surfaces to the static aerodynamic characteristics at Mach numbers of 0.8, 1.3, and 1.75 and Reynolds number of 625,000 based on body diameter. Data were obtained at angles of attack ranging from 0 deg to 24 deg for various stages of model build-up (i.e., with and without canard and/or tail surfaces). In addition, two different sets of canards and tail surfaces were investigated. For the canard and tail arrangements investigated, the model was trimmable at angles of attack up to about 7 deg with canard deflections of about 10 deg. Also, the tail arrangements studied provided ample pitch stability.

Study of Swept Angle Effects on Grid Fins Aerodynamics Performance

NASA Astrophysics Data System (ADS)

Faza, G. A.; Fadillah, H.; Silitonga, F. Y.; Agoes Moelyadi, Mochamad

2018-04-01

Grid fin is an aerodynamic control surface that usually used on missiles and rockets. In the recent several years many researches have conducted to develop a more efficient grid fins. There are many possibilities of geometric combination could be done to improve aerodynamics characteristic of a grid fin. This paper will only discuss about the aerodynamics characteristics of grid fins compared by another grid fins with different swept angle. The methodology that used to compare the aerodynamics is Computational Fluid Dynamics (CFD). The result of this paper might be used for future studies to answer our former question or as a reference for related studies.

Evaluation of Turbulence-Model Performance in Jet Flows

NASA Technical Reports Server (NTRS)

Woodruff, S. L.; Seiner, J. M.; Hussaini, M. Y.; Erlebacher, G.

2001-01-01

The importance of reducing jet noise in both commercial and military aircraft applications has made jet acoustics a significant area of research. A technique for jet noise prediction commonly employed in practice is the MGB approach, based on the Lighthill acoustic analogy. This technique requires as aerodynamic input mean flow quantities and turbulence quantities like the kinetic energy and the dissipation. The purpose of the present paper is to assess existing capabilities for predicting these aerodynamic inputs. Two modern Navier-Stokes flow solvers, coupled with several modern turbulence models, are evaluated by comparison with experiment for their ability to predict mean flow properties in a supersonic jet plume. Potential weaknesses are identified for further investigation. Another comparison with similar intent is discussed by Barber et al. The ultimate goal of this research is to develop a reliable flow solver applicable to the low-noise, propulsion-efficient, nozzle exhaust systems being developed in NASA focused programs. These programs address a broad range of complex nozzle geometries operating in high temperature, compressible, flows. Seiner et al. previously discussed the jet configuration examined here. This convergent-divergent nozzle with an exit diameter of 3.6 inches was designed for an exhaust Mach number of 2.0 and a total temperature of 1680 F. The acoustic and aerodynamic data reported by Seiner et al. covered a range of jet total temperatures from 104 F to 2200 F at the fully-expanded nozzle pressure ratio. The aerodynamic data included centerline mean velocity and total temperature profiles. Computations were performed independently with two computational fluid dynamics (CFD) codes, ISAAC and PAB3D. Turbulence models employed include the k-epsilon model, the Gatski-Speziale algebraic-stress model and the Girimaji model, with and without the Sarkar compressibility correction. Centerline values of mean velocity and mean temperature are compared with experimental data.

The Mars Exploration Rovers Entry Descent and Landing and the Use of Aerodynamic Decelerators

NASA Technical Reports Server (NTRS)

Steltzner, Adam; Desai, Prasun; Lee, Wayne; Bruno, Robin

2003-01-01

The Mars Exploration Rovers (MER) project, the next United States mission to the surface of Mars, uses aerodynamic decelerators in during its entry, descent and landing (EDL) phase. These two identical missions (MER-A and MER-B), which deliver NASA s largest mobile science suite to date to the surface of Mars, employ hypersonic entry with an ablative energy dissipating aeroshell, a supersonic/subsonic disk-gap-band parachute and an airbag landing system within EDL. This paper gives an overview of the MER EDL system and speaks to some of the challenges faced by the various aerodynamic decelerators.

Research on unsteady transonic flow theory

NASA Technical Reports Server (NTRS)

Revell, J. D.

1973-01-01

A two-dimensional theory is considered for the unsteady flow disturbances caused by aeroelastic deformations of a thick wing at high subsonic freestream Mach numbers, having a single, internally embedded supercritical (locally supersonic) steady flow region adjacent to the low pressure side of the wing. The theory develops a matrix of unsteady aerodynamic influence coefficients (AICs) suitable as a strip theory for aeroelastic analysis of large aspect ratio thick wings of moderate sweep, typical of a wide class of current and future aircraft. The theory derives the linearized unsteady flow solutions separately for both the subcritical and supercritical regions. These solutions are coupled together to give the requisite (wing pressure-downwash) AICs by the intermediate step of defining flow disturbances on the sonic line, and at the shock wave; these intermediate quantities are then algebraically eliminated by expressing them in terms of the wing surface downwash.

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.

High Speed Aerodynamic Characteristics of the GAF0PH Aerofoil

DTIC Science & Technology

1980-09-01

upper surface of the aerofoil for angles of incidence greater than 210. POSTAL ADDRESS: Chief Superintendent, Aeronautical Research Laboratories, Box...kCLAERO-.NOTE3 98 -AR-002-223 -LEVEL m DEPARTMENT OF DEFENCE 00 DEFENCE SCIENCE AND TECHNOLOGY ORGANISATION AERONAUTICAL RESEARCH LABORATORIES...MELBOURNE, VICTORIA AERODYNAMICS NOTE 398 ’,\\ HIGH SPEED AERODYNAMIC CHARACTERISTICS OF THE GAFPH AEROFOIL by ~B D :, . , .IR-© Approved for Public Release

X-33 Computational Aeroheating/Aerodynamic Predictions and Comparisons With Experimental Data

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Thompson, Richard A.; Berry, Scott A.; Horvath, Thomas J.; Murphy, Kelly J.; Nowak, Robert J.; Alter, Stephen J.

2003-01-01

This report details a computational fluid dynamics study conducted in support of the phase II development of the X-33 vehicle. Aerodynamic and aeroheating predictions were generated for the X-33 vehicle at both flight and wind-tunnel test conditions using two finite-volume, Navier-Stokes solvers. Aerodynamic computations were performed at Mach 6 and Mach 10 wind-tunnel conditions for angles of attack from 10 to 50 with body-flap deflections of 0 to 20. Additional aerodynamic computations were performed over a parametric range of free-stream conditions at Mach numbers of 4 to 10 and angles of attack from 10 to 50. Laminar and turbulent wind-tunnel aeroheating computations were performed at Mach 6 for angles of attack of 20 to 40 with body-flap deflections of 0 to 20. Aeroheating computations were performed at four flight conditions with Mach numbers of 6.6 to 8.9 and angles of attack of 10 to 40. Surface heating and pressure distributions, surface streamlines, flow field information, and aerodynamic coefficients from these computations are presented, and comparisons are made with wind-tunnel data.

Effects of Nose Radius and Aerodynamic Loading on Leading Edge Receptivity

NASA Technical Reports Server (NTRS)

Hammerton, P. W.; Kerschen, E. J.

1998-01-01

An analysis is presented of the effects of airfoil thickness and mean aerodynamic loading on boundary-layer receptivity in the leading-edge region. The case of acoustic free-stream disturbances, incident on a thin cambered airfoil with a parabolic leading edge in a low Mach number flow, is considered. An asymptotic analysis based on large Reynolds number is developed, supplemented by numerical results. The airfoil thickness distribution enters the theory through a Strouhal number based on the nose radius of the airfoil, S = (omega)tau(sub n)/U, where omega is the frequency of the acoustic wave and U is the mean flow speed. The influence of mean aerodynamic loading enters through an effective angle-of-attack parameter ti, related to flow around the leading edge from the lower surface to the upper. The variation of the receptivity level is analyzed as a function of S, mu, and characteristics of the free-stream acoustic wave. For an unloaded leading edge, a finite nose radius dramatically reduces the receptivity level compared to that for a flat plate, the amplitude of the instability waves in the boundary layer being decreased by an order of magnitude when S = 0.3. Modest levels of aerodynamic loading are found to further decrease the receptivity level for the upper surface of the airfoil, while an increase in receptivity level occurs for the lower surface. For larger angles of attack close to the critical angle for boundary layer separation, a local rise in the receptivity level occurs for the upper surface, while for the lower surface the receptivity decreases. The effects of aerodynamic loading are more pronounced at larger values of S. Oblique acoustic waves produce much higher receptivity levels than acoustic waves propagating downstream parallel to the airfoil chord.

Design and Execution of the Hypersonic Inflatable Aerodynamic Decelerator Large-Article Wind Tunnel Experiment

NASA Technical Reports Server (NTRS)

Cassell, Alan M.

2013-01-01

The testing of 3- and 6-meter diameter Hypersonic Inflatable Aerodynamic Decelerator (HIAD) test articles was completed in the National Full-Scale Aerodynamics Complex 40 ft x 80 ft Wind Tunnel test section. Both models were stacked tori, constructed as 60 degree half-angle sphere cones. The 3-meter HIAD was tested in two configurations. The first 3-meter configuration utilized an instrumented flexible aerodynamic skin covering the inflatable aeroshell surface, while the second configuration employed a flight-like flexible thermal protection system. The 6-meter HIAD was tested in two structural configurations (with and without an aft-mounted stiffening torus near the shoulder), both utilizing an instrumented aerodynamic skin.

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.

A program to evaluate a control system based on feedback of aerodynamic pressure differentials

NASA Technical Reports Server (NTRS)

Levy, D. W.; Finn, P.; Roskam, J.

1981-01-01

The use of aerodynamic pressure differentials to position a control surface is evaluated. The system is a differential pressure command loop, analogous to a position command loop, where the surface is commanded to move until a desired differential pressure across the surface is achieved. This type of control is more direct and accurate because it is the differential pressure which causes the control forces and moments. A frequency response test was performed in a low speed wind tunnel to measure the performance of the system. Both pressure and position feedback were tested. The pressure feedback performed as well as position feedback implying that the actuator, with a break frequency on the order of 10 Rad/sec, was the limiting component. Theoretical considerations indicate that aerodynamic lags will not appear below frequencies of 50 Rad/sec, or higher.

The influence of flight style on the aerodynamic properties of avian wings as fixed lifting surfaces

PubMed Central

Dimitriadis, Grigorios; Nudds, Robert L.

2016-01-01

The diversity of wing morphologies in birds reflects their variety of flight styles and the associated aerodynamic and inertial requirements. Although the aerodynamics underlying wing morphology can be informed by aeronautical research, important differences exist between planes and birds. In particular, birds operate at lower, transitional Reynolds numbers than do most aircraft. To date, few quantitative studies have investigated the aerodynamic performance of avian wings as fixed lifting surfaces and none have focused upon the differences between wings from different flight style groups. Dried wings from 10 bird species representing three distinct flight style groups were mounted on a force/torque sensor within a wind tunnel in order to test the hypothesis that wing morphologies associated with different flight styles exhibit different aerodynamic properties. Morphological differences manifested primarily as differences in drag rather than lift. Maximum lift coefficients did not differ between groups, whereas minimum drag coefficients were lowest in undulating flyers (Corvids). The lift to drag ratios were lower than in conventional aerofoils and data from free-flying soaring species; particularly in high frequency, flapping flyers (Anseriformes), which do not rely heavily on glide performance. The results illustrate important aerodynamic differences between the wings of different flight style groups that cannot be explained solely by simple wing-shape measures. Taken at face value, the results also suggest that wing-shape is linked principally to changes in aerodynamic drag, but, of course, it is aerodynamics during flapping and not gliding that is likely to be the primary driver. PMID:27781155

Aerodynamic Control-Augmentation Devices For Saturn-Class Launch Vehicles With Aft Centers Of Gravity

NASA Technical Reports Server (NTRS)

Barret, Chris

1995-01-01

Report describes study of aerodynamic flight-control-augmentation devices proposed for use in increasing payload capabilities of future launch vehicles by allowing more aft centers of gravity. Proposed all-movable devices not only provide increased control authority during ascent trajectory, but also reduce engine gimballing requirements and enhance crew safety. Report proposes various aerodynamic control surfaces mounted fore and aft on Saturn-class launch vehicle.

Aerodynamic Shutoff Valve

NASA Technical Reports Server (NTRS)

Horstman, Raymond H.

1992-01-01

Aerodynamic flow achieved by adding fixed fairings to butterfly valve. When valve fully open, fairings align with butterfly and reduce wake. Butterfly free to turn, so valve can be closed, while fairings remain fixed. Design reduces turbulence in flow of air in internal suction system. Valve aids in development of improved porous-surface boundary-layer control system to reduce aerodynamic drag. Applications primarily aerospace. System adapted to boundary-layer control on high-speed land vehicles.

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.

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.

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.

Three-dimensional aerodynamic shape optimization of supersonic delta wings

NASA Technical Reports Server (NTRS)

Burgreen, Greg W.; Baysal, Oktay

1994-01-01

A recently developed three-dimensional aerodynamic shape optimization procedure AeSOP(sub 3D) is described. This procedure incorporates some of the most promising concepts from the area of computational aerodynamic analysis and design, specifically, discrete sensitivity analysis, a fully implicit 3D Computational Fluid Dynamics (CFD) methodology, and 3D Bezier-Bernstein surface parameterizations. The new procedure is demonstrated in the preliminary design of supersonic delta wings. Starting from a symmetric clipped delta wing geometry, a Mach 1.62 asymmetric delta wing and two Mach 1. 5 cranked delta wings were designed subject to various aerodynamic and geometric constraints.

Advanced propeller aerodynamic analysis

NASA Technical Reports Server (NTRS)

Bober, L. J.

1980-01-01

The analytical approaches as well as the capabilities of three advanced analyses for predicting propeller aerodynamic performance are presented. It is shown that two of these analyses use a lifting line representation for the propeller blades, and the third uses a lifting surface representation.

Abatement of an aircraft exhaust plume using aerodynamic baffles.

PubMed

Bennett, Michael; Christie, Simon M; Graham, Angus; Garry, Kevin P; Velikov, Stefan; Poll, D Ian; Smith, Malcolm G; Mead, M Iqbal; Popoola, Olalekan A M; Stewart, Gregor B; Jones, Roderic L

2013-03-05

The exhaust jet from a departing commercial aircraft will eventually rise buoyantly away from the ground; given the high thrust/power (i.e., momentum/buoyancy) ratio of modern aero-engines, however, this is a slow process, perhaps requiring ∼ 1 min or more. Supported by theoretical and wind tunnel modeling, we have experimented with an array of aerodynamic baffles on the surface behind a set of turbofan engines of 124 kN thrust. Lidar and point sampler measurements show that, as long as the intervention takes place within the zone where the Coanda effect holds the jet to the surface (i.e., within about 70 m in this case), then quite modest surface-mounted baffles can rapidly lift the jet away from the ground. This is of potential benefit in abating both surface concentrations and jet blast downstream. There is also some modest acoustic benefit. By distributing the aerodynamic lift and drag across an array of baffles, each need only be a fraction of the height of a single blast fence.

Launch vehicle flight control augmentation using smart materials and advanced composites (CDDF Project 93-05)

NASA Technical Reports Server (NTRS)

Barret, C.

1995-01-01

The Marshall Space Flight Center has a rich heritage of launch vehicles that have used aerodynamic surfaces for flight stability such as the Saturn vehicles and flight control such as on the Redstone. Recently, due to aft center-of-gravity locations on launch vehicles currently being studied, the need has arisen for the vehicle control augmentation that is provided by these flight controls. Aerodynamic flight control can also reduce engine gimbaling requirements, provide actuator failure protection, enhance crew safety, and increase vehicle reliability, and payload capability. In the Saturn era, NASA went to the Moon with 300 sq ft of aerodynamic surfaces on the Saturn V. Since those days, the wealth of smart materials and advanced composites that have been developed allow for the design of very lightweight, strong, and innovative launch vehicle flight control surfaces. This paper presents an overview of the advanced composites and smart materials that are directly applicable to launch vehicle control surfaces.

Mechanical power curve measured in the wake of pied flycatchers indicates modulation of parasite power across flight speeds.

PubMed

Johansson, L Christoffer; Maeda, Masateru; Henningsson, Per; Hedenström, Anders

2018-01-01

How aerodynamic power required for animal flight varies with flight speed determines optimal speeds during foraging and migratory flight. Despite its relevance, aerodynamic power provides an elusive quantity to measure directly in animal flight. Here, we determine the aerodynamic power from wake velocity fields, measured using tomographical particle image velocimetry, of pied flycatchers flying freely in a wind tunnel. We find a shallow U-shaped power curve, which is flatter than expected by theory. Based on how the birds vary body angle with speed, we speculate that the shallow curve results from increased body drag coefficient and body frontal area at lower flight speeds. Including modulation of body drag in the model results in a more reasonable fit with data than the traditional model. From the wake structure, we also find a single starting vortex generated from the two wings during the downstroke across flight speeds (1-9 m s -1 ). This is accomplished by the arm wings interacting at the beginning of the downstroke, generating a unified starting vortex above the body of the bird. We interpret this as a mechanism resulting in a rather uniform downwash and low induced power, which can help explain the higher aerodynamic performance in birds compared with bats. © 2018 The Author(s).

Aerodynamic characteristics of horizontal tail surfaces

NASA Technical Reports Server (NTRS)

Silverstein, Abe; Katzoff, S

1940-01-01

Collected data are presented on the aerodynamic characteristics of 17 horizontal tail surfaces including several with balanced elevators and two with end plates. Curves are given for coefficients of normal force, drag, and elevator hinge moment. A limited analysis of the results has been made. The normal-force coefficients are in better agreement with the lifting-surface theory of Prandtl and Blenk for airfoils of low aspect ratio than with the usual lifting-line theory. Only partial agreement exists between the elevator hinge-moment coefficients and those predicted by Glauert's thin-airfoil theory.

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.

Transonic aerodynamic characteristics of a proposed wing-body reusable launch vehicle concept

NASA Technical Reports Server (NTRS)

Springer, A. M.

1995-01-01

A proposed wing-body reusable launch vehicle was tested in the NASA Marshall Space Flight Center's 14 x 14-inch trisonic wind tunnel during the winter of 1994. This test resulted in the vehicle's subsonic and transonic, Mach 0.3 to 1.96, longitudinal and lateral aerodynamic characteristics. The effects of control surface deflections on the basic vehicle's aerodynamics, including a body flap, elevons, ailerons, and tip fins, are presented.

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.

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.

Endplate effect on aerodynamic characteristics of threedimensional wings in close free surface proximity

NASA Astrophysics Data System (ADS)

Jung, Jae Hwan; Kim, Mi Jeong; Yoon, Hyun Sik; Hung, Pham Anh; Chun, Ho Hwan; Park, Dong Woo

2012-12-01

We investigated the aerodynamic characteristics of a three-dimensional (3D) wing with an endplate in the vicinity of the free surface by solving incompressible Navier-Stokes equations with the turbulence closure model. The endplate causes a blockage effect on the flow, and an additional viscous effect especially near the endplate. These combined effects of the endplate significantly reduce the magnitudes of the velocities under the lower surface of the wing, thereby enhancing aerodynamic performance in terms of the force coefficients. The maximum lift-to-drag ratio of a wing with an endplate is increased 46% compared to that of wing without an endplate at the lowest clearance. The tip vortex of a wing-with-endplate (WWE) moved laterally to a greater extent than that of a wing-without-endplate (WOE). This causes a decrease in the induced drag, resulting in a reduction in the total drag.

Experimental aeroelastic control using adaptive wing model concepts

NASA Astrophysics Data System (ADS)

Costa, Antonio P.; Moniz, Paulo A.; Suleman, Afzal

2001-06-01

The focus of this study is to evaluate the aeroelastic performance and control of adaptive wings. Ailerons and flaps have been designed and implemented into 3D wings for comparison with adaptive structures and active aerodynamic surface control methods. The adaptive structures concept, the experimental setup and the control design are presented. The wind-tunnel tests of the wing models are presented for the open- and closed-loop systems. The wind tunnel testing has allowed for quantifying the effectiveness of the piezoelectric vibration control of the wings, and also provided performance data for comparison with conventional aerodynamic control surfaces. The results indicate that a wing utilizing skins as active structural elements with embedded piezoelectric actuators can be effectively used to improve the aeroelastic response of aeronautical components. It was also observed that the control authority of adaptive wings is much greater than wings using conventional aerodynamic control surfaces.

Aerodynamic characteristics of a canard-controlled missile at Mach numbers of 1.5 and 2.0.

NASA Technical Reports Server (NTRS)

Kassner, D. L.; Wettlaufer, B.

1977-01-01

A typical missile model with nose mounted canards and cruciform tail surfaces was tested in the Ames 6- by 6-Foot Wind Tunnel to determine the contributions of the component aerodynamic surfaces to the static aerodynamic characteristics at Mach numbers of 1.5 and 2.0 and Reynolds number of 1 million based on body diameter. Data were obtained at angles of attack ranging from -3 deg to 12 deg for various stages of model build-up (i.e., with and without canard and/or tail surfaces). Results were obtained both with the model unrolled and rolled 45 deg. For the canard and tail arrangements investigated, the model was trimmable at angles of attack up to about 10 deg with canard deflections of 9 deg. Also, the tail arrangements studied provided ample pitch stability. there were no appreciable effects of model roll orientation.

Interactive Graphics Analysis for Aircraft Design

NASA Technical Reports Server (NTRS)

Townsend, J. C.

1983-01-01

Program uses higher-order far field drag minimization. Computer program WDES WDEM preliminary aerodynamic design tool for one or two interacting, subsonic lifting surfaces. Subcritical wing design code employs higher-order far-field drag minimization technique. Linearized aerodynamic theory used. Program written in FORTRAN IV.

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.

Testing of the Trim Tab Parametric Model in NASA Langley's Unitary Plan Wind Tunnel

NASA Technical Reports Server (NTRS)

Murphy, Kelly J.; Watkins, Anthony N.; Korzun, Ashley M.; Edquist, Karl T.

2013-01-01

In support of NASA's Entry, Descent, and Landing technology development efforts, testing of Langley's Trim Tab Parametric Models was conducted in Test Section 2 of NASA Langley's Unitary Plan Wind Tunnel. The objectives of these tests were to generate quantitative aerodynamic data and qualitative surface pressure data for experimental and computational validation and aerodynamic database development. Six component force-and-moment data were measured on 38 unique, blunt body trim tab configurations at Mach numbers of 2.5, 3.5, and 4.5, angles of attack from -4deg to +20deg, and angles of sideslip from 0deg to +8deg. Configuration parameters investigated in this study were forebody shape, tab area, tab cant angle, and tab aspect ratio. Pressure Sensitive Paint was used to provide qualitative surface pressure mapping for a subset of these flow and configuration variables. Over the range of parameters tested, the effects of varying tab area and tab cant angle were found to be much more significant than varying tab aspect ratio relative to key aerodynamic performance requirements. Qualitative surface pressure data supported the integrated aerodynamic data and provided information to aid in future analyses of localized phenomena for trim tab configurations.

Film cooling effectiveness on a large angle blunt cone flying at hypersonic speed

NASA Astrophysics Data System (ADS)

Sahoo, Niranjan; Kulkarni, Vinayak; Saravanan, S.; Jagadeesh, G.; Reddy, K. P. J.

2005-03-01

Effectiveness of film cooling technique to reduce convective heating rates for a large angle blunt cone flying at hypersonic Mach number and its effect on the aerodynamic characteristics is investigated experimentally by measuring surface heat-transfer rates and aerodynamic drag coefficient simultaneously. The test model is a 60° apex-angle blunt cone with an internally mounted accelerometer balance system for measuring aerodynamic drag and an array of surface mounted platinum thin film gauges for measuring heat-transfer rates. The coolant gas (air, carbon dioxide, and/or helium) is injected into the hypersonic flow at the nose of the test model. The experiments are performed at a flow free stream Mach number of 5.75 and 0° angle of attack for stagnation enthalpies of 1.16MJ/kg and 1.6MJ/kg with and without gas injection. About 30%-45% overall reduction in heat-transfer rates is observed with helium as coolant gas except at stagnation regions. With all other coolants, the reduction in surface heat-transfer rate is between 10%-25%. The aerodynamic drag coefficient is found to increase by 12% with helium injection whereas with other gases this increase is about 27%.

Numerical Study of Steady and Unsteady Canard-Wing-Body Aerodynamics

NASA Technical Reports Server (NTRS)

Eugene, L. Tu

1996-01-01

The use of canards in advanced aircraft for control and improved aerodynamic performance is a topic of continued interest and research. In addition to providing maneuver control and trim, the influence of canards on wing aerodynamics can often result in increased maximum lift and decreased trim drag. In many canard-configured aircraft, the main benefits of canards are realized during maneuver or other dynamic conditions. Therefore, the detailed study and understanding of canards requires the accurate prediction of the non-linear unsteady aerodynamics of such configurations. For close-coupled canards, the unsteady aerodynamic performance associated with the canard-wing interaction is of particular interest. The presence of a canard in close proximity to the wing results in a highly coupled canard-wing aerodynamic flowfield which can include downwash/upwash effects, vortex-vortex interactions and vortex-surface interactions. For unsteady conditions, these complexities of the canard-wing flowfield are further increased. The development and integration of advanced computational technologies provide for the time-accurate Navier-Stokes simulations of the steady and unsteady canard-wing-body flox,fields. Simulation, are performed for non-linear flight regimes at transonic Mach numbers and for a wide range of angles of attack. For the static configurations, the effects of canard positioning and fixed deflection angles on aerodynamic performance and canard-wing vortex interaction are considered. For non-static configurations, the analyses of the canard-wing body flowfield includes the unsteady aerodynamics associated with pitch-up ramp and pitch oscillatory motions of the entire geometry. The unsteady flowfield associated with moving canards which are typically used as primary control surfaces are considered as well. The steady and unsteady effects of the canard on surface pressure integrated forces and moments, and canard-wing vortex interaction are presented in detail including the effects of the canard on the static and dynamic stability characteristics. The current study provides an understanding of the steady and unsteady canard-wing-body flowfield. Emphasis is placed on the effects of the canard on aerodynamic performance as well as the detailed flow physics of the canard-wing flowfield interactions. The computational tools developed to accurately predict the time-accurate flowfield of moving canards provides for the capability of coupled fluids-controls simulations desired in the detailed design and analysis of advanced aircraft.

A flight experiment to measure rarefied-flow aerodynamics

NASA Technical Reports Server (NTRS)

Blanchard, Robert C.

1990-01-01

A flight experiment to measure rarefied-flow aerodynamics of a blunt lifting body is being developed by NASA. This experiment, called the Rarefied-Flow Aerodynamic Measurement Experiment (RAME), is part of the Aeroassist Flight Experiment (AFE) mission, which is a Pathfinder design tool for aeroassisted orbital transfer vehicles. The RAME will use flight measurements from accelerometers, rate gyros, and pressure transducers, combined with knowledge of AFE in-flight mass properties and trajectory, to infer aerodynamic forces and moments in the rarefied-flow environment, including transition into the hypersonic continuum regime. Preflight estimates of the aerodynamic measurements are based upon environment models, existing computer simulations, and ground test results. Planned maneuvers at several altitudes will provide a first-time opportunity to examine gas-surface accommondation effects on aerodynamic coefficients in an environment of changing atmospheric composition. A description is given of the RAME equipment design.

Hypersonic aerospace vehicle leading edge cooling using heat pipe, transpiration and film cooling techniques

NASA Astrophysics Data System (ADS)

Modlin, James Michael

An investigation was conducted to study the feasibility of cooling hypersonic vehicle leading edge structures exposed to severe aerodynamic surface heat fluxes using a combination of liquid metal heat pipes and surface mass transfer cooling techniques. A generalized, transient, finite difference based hypersonic leading edge cooling model was developed that incorporated these effects and was demonstrated on an assumed aerospace plane-type wing leading edge section and a SCRAMJET engine inlet leading edge section. The hypersonic leading edge cooling model was developed using an existing, experimentally verified heat pipe model. Two applications of the hypersonic leading edge cooling model were examined. An assumed aerospace plane-type wing leading edge section exposed to a severe laminar, hypersonic aerodynamic surface heat flux was studied. A second application of the hypersonic leading edge cooling model was conducted on an assumed one-quarter inch nose diameter SCRAMJET engine inlet leading edge section exposed to both a transient laminar, hypersonic aerodynamic surface heat flux and a type 4 shock interference surface heat flux. The investigation led to the conclusion that cooling leading edge structures exposed to severe hypersonic flight environments using a combination of liquid metal heat pipe, surface transpiration, and film cooling methods appeared feasible.

Overview of the Cranked-Arrow Wing Aerodynamics Project International

NASA Technical Reports Server (NTRS)

Obara, Clifford J.; Lamar, John E.

2008-01-01

This paper provides a brief history of the F-16XL-1 aircraft, its role in the High Speed Research program and how it was morphed into the Cranked Arrow Wing Aerodynamics Project. Various flight, wind-tunnel and Computational Fluid Dynamics data sets were generated as part of the project. These unique and open flight datasets for surface pressures, boundary-layer profiles and skin-friction distributions, along with surface flow data, are described and sample data comparisons given. This is followed by a description of how the project became internationalized to be known as Cranked Arrow Wing Aerodynamics Project International and is concluded by an introduction to the results of a four year computational predictive study of data collected at flight conditions by participating researchers.

Application of the ASP3D Computer Program to Unsteady Aerodynamic and Aeroelastic Analyses

NASA Technical Reports Server (NTRS)

Batina, John T.

2006-01-01

A new computer program has been developed called ASP3D (Advanced Small Perturbation - 3D), which solves the small perturbation potential flow equation in an advanced form including mass-consistent surface and trailing wake boundary conditions, and entropy, vorticity, and viscous effects. The purpose of the program is for unsteady aerodynamic and aeroelastic analyses, especially in the nonlinear transonic flight regime. The program exploits the simplicity of stationary Cartesian meshes with the movement or deformation of the configuration under consideration incorporated into the solution algorithm through a planar surface boundary condition. The paper presents unsteady aerodynamic and aeroelastic applications of ASP3D to assess the time dependent capability and demonstrate various features of the code.

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.

Ascent Aerodynamic Pressure Distributions on WB001

NASA Technical Reports Server (NTRS)

Vu, B.; Ruf, J.; Canabal, F.; Brunty, J.

1996-01-01

To support the reusable launch vehicle concept study, the aerodynamic data and surface pressure for WB001 were predicted using three computational fluid dynamic (CFD) codes at several flow conditions between code to code and code to aerodynamic database as well as available experimental data. A set of particular solutions have been selected and recommended for use in preliminary conceptual designs. These computational fluid dynamic (CFD) results have also been provided to the structure group for wing loading analysis.

Application of the FADS system on the Re-entry Module

NASA Astrophysics Data System (ADS)

Zhen, Huang

2016-07-01

The aerodynamic model for Flush Air Data Sensing System (FADS) is built based on the surface pressure distribution obtained through the pressure orifices laid on specific positions of the surface,and the flight parameters,such as angle of attack,angle of side-slip,Mach number,free-stream static pressure and dynamic pressure are inferred from the aerodynamic model.The flush air data sensing system (FADS) has been used on several flight tests of aircraft and re-entry vehicle,such as,X-15,space shuttle,F-14,X-33,X-43A and so on. This paper discusses the application of the FADS on the re-entry module with blunt body to obtain high-precision aerodynamic parameters.First of all,a basic theory and operating principle of the FADS is shown.Then,the applications of the FADS on typical aircrafts and re-entry vehicles are described.Thirdly,the application mode on the re-entry module with blunt body is discussed in detail,including aerodynamic simulation,pressure distribution,trajectory reconstruction and the hardware shoule be used,such as flush air data sensing system(FADS),inertial navigation system (INS),data acquisition system,data storage system.Finally,ablunt module re-entry flight test from low earth orbit (LEO) is planned to obtain aerodynamic parameters and amend the aerodynamic model with this FADS system data.The results show that FADS system can be applied widely in re-entry module with blunt bodies.

Overview of Supersonic Aerodynamics Measurement Techniques in the NASA Langley Unitary Plan Wind Tunnel

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

2007-01-01

An overview is given of selected measurement techniques used in the NASA Langley Research Center (NASA LaRC) Unitary Plan Wind Tunnel (UPWT) to determine the aerodynamic characteristics of aerospace vehicles operating at supersonic speeds. A broad definition of a measurement technique is adopted in this paper and is any qualitative or quantitative experimental approach that provides information leading to the improved understanding of the supersonic aerodynamic characteristics. On-surface and off-surface measurement techniques used to obtain discrete (point) and global (field) measurements and planar and global flow visualizations are described, and examples of all methods are included. The discussion is limited to recent experiences in the UPWT and is, therefore, not an exhaustive review of existing experimental techniques. The diversity and high quality of the measurement techniques and the resultant data illustrate the capabilities of a ground-based experimental facility and the key role that it plays in the advancement of our understanding, prediction, and control of supersonic aerodynamics.

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.

Aerodynamic Characteristics of Two Waverider-Derived Hypersonic Cruise Configurations

NASA Technical Reports Server (NTRS)

Cockrell, Charles E., Jr.; Huebner, Lawrence D.; Finley, Dennis B.

1996-01-01

An evaluation was made on the effects of integrating the required aircraft components with hypersonic high-lift configurations known as waveriders to create hypersonic cruise vehicles. Previous studies suggest that waveriders offer advantages in aerodynamic performance and propulsion/airframe integration (PAI) characteristics over conventional non-waverider hypersonic shapes. A wind-tunnel model was developed that integrates vehicle components, including canopies, engine components, and control surfaces, with two pure waverider shapes, both conical-flow-derived waveriders for a design Mach number of 4.0. Experimental data and limited computational fluid dynamics (CFD) solutions were obtained over a Mach number range of 1.6 to 4.63. The experimental data show the component build-up effects and the aerodynamic characteristics of the fully integrated configurations, including control surface effectiveness. The aerodynamic performance of the fully integrated configurations is not comparable to that of the pure waverider shapes, but is comparable to previously tested hypersonic models. Both configurations exhibit good lateral-directional stability characteristics.

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.

USB noise reduction by nozzle and flap modifications

NASA Technical Reports Server (NTRS)

Hayden, R. E.

1976-01-01

The development of concepts for reducing upper surface blown flap noise at the source through flap modifications and special nozzles is reviewed. In particular, recent results obtained on the aerodynamic and acoustic performance of flaps with porous surfaces near the trailing edge and multi-slotted nozzles are reviewed. Considerable reduction (6-10 db) of the characteristic low frequency peak is shown. The aerodynamic performance is compared with conventional systems, and prospects for future improvements are discussed.

Estimation of Supersonic Stage Separation Aerodynamics of Winged-Body Launch Vehicles Using Response Surface Methods

NASA Technical Reports Server (NTRS)

Erickson, Gary E.; Deloach, Richard

2008-01-01

A collection of statistical and mathematical techniques referred to as response surface methodology was used to estimate the longitudinal stage separation aerodynamic characteristics of a generic, bimese, winged multi-stage launch vehicle configuration using data obtained on small-scale models at supersonic speeds in the NASA Langley Research Center Unitary Plan Wind Tunnel. The simulated Mach 3 staging was dominated by multiple shock wave interactions between the orbiter and booster vehicles throughout the relative spatial locations of interest. This motivated a partitioning of the overall inference space into several contiguous regions within which the separation aerodynamics were presumed to be well-behaved and estimable using cuboidal and spherical central composite designs capable of fitting full second-order response functions. The primary goal was to approximate the underlying overall aerodynamic response surfaces of the booster vehicle in belly-to-belly proximity to the orbiter vehicle using relatively simple, lower-order polynomial functions that were piecewise-continuous across the full independent variable ranges of interest. The quality of fit and prediction capabilities of the empirical models were assessed in detail, and the issue of subspace boundary discontinuities was addressed. The potential benefits of augmenting the central composite designs to full third order using computer-generated D-optimality criteria were also evaluated. The usefulness of central composite designs, the subspace sizing, and the practicality of fitting low-order response functions over a partitioned inference space dominated by highly nonlinear and possibly discontinuous shock-induced aerodynamics are discussed.

Space-Time Interface-Tracking Computations with Contact Between Solid Surfaces

DTIC Science & Technology

2014-04-01

parachute FSI [70, 72, 73, 75, 53, 55, 46, 51, 57], flapping-wing aerodynamics [48, 50], and wind - turbine rotor and tower aerodynamics [61]. It can...48, 50], and wind - turbine rotor and tower aerodynamics with the blades passing the tower 18 4 close [61]. As mentioned in [16], one of course...9] Y. Bazilevs, M.-C. Hsu, I. Akkerman, S. Wright, K. Takizawa, B. Henicke, T. Spielman, and T. E. Tezduyar. 3D simulation of wind turbine rotors at

A Feasibility Study on the Control of a Generic Air Vehicle Using Control Moment Gyros

NASA Technical Reports Server (NTRS)

Lim, Kyong B.; Moerder, Daniel D.

2006-01-01

This paper examines feasibility and performance issues in using Control Moment Gyroscopes (CMGs) to control the attitude of a fixed-wing aircraft. The paper describes a control system structure that permits allocating control authority and bandwidth between a CMG system and conventional aerodynamic control surfaces to stabilize a vehicle with neutral aerodynamic stability. A simulation study explores the interplay between aerodynamic and CMG effects, and indicates desirable physical characteristics for a CMG system to be used for aircraft attitude control.

Nonlinear aerodynamic wing design

NASA Technical Reports Server (NTRS)

Bonner, Ellwood

1985-01-01

The applicability of new nonlinear theoretical techniques is demonstrated for supersonic wing design. The new technology was utilized to define outboard panels for an existing advanced tactical fighter model. Mach 1.6 maneuver point design and multi-operating point compromise surfaces were developed and tested. High aerodynamic efficiency was achieved at the design conditions. A corollary result was that only modest supersonic penalties were incurred to meet multiple aerodynamic requirements. The nonlinear potential analysis of a practical configuration arrangement correlated well with experimental data.

Flap effectiveness appraisal for winged re-entry vehicles

NASA Astrophysics Data System (ADS)

de Rosa, Donato; Pezzella, Giuseppe; Donelli, Raffaele S.; Viviani, Antonio

2016-05-01

The interactions between shock waves and boundary layer are commonplace in hypersonic aerodynamics. They represent a very challenging design issue for hypersonic vehicle. A typical example of shock wave boundary layer interaction is the flowfield past aerodynamic surfaces during control. As a consequence, such flow interaction phenomena influence both vehicle aerodynamics and aerothermodynamics. In this framework, the present research effort describes the numerical activity performed to simulate the flowfield past a deflected flap in hypersonic flowfield conditions for a winged re-entry vehicle.

Shear-Sensitive Liquid Crystal Coating Method Applied Through Transparent Test Surfaces

NASA Technical Reports Server (NTRS)

Reda, Daniel C.; Wilder, Michael C.

1999-01-01

Research conducted at NASA Ames Research Center has shown that the color-change response of a shear-sensitive liquid crystal coating (SSLCC) to aerodynamic shear depends on both the magnitude of the local shear vector and its direction relative to the observer's in-plane line of sight. In conventional applications, the surface of the SSLCC exposed to aerodynamic shear is illuminated with white light from the normal direction and observed from an oblique above-plane view angle of order 30 deg. In this top-light/top-view mode, shear vectors with components directed away from the observer cause the SSLCC to exhibit color-change responses. At any surface point, the maximum color change (measured from the no-shear red or orange color) always occurs when the local vector is aligned with, and directed away from, the observer. The magnitude of the color change at this vector-observer-aligned orientation scales directly with shear stress magnitude. Conversely, any surface point exposed to a shear vector with a component directed toward the observer exhibits a non-color-change response, always characterized by a rusty-red or brown color, independent of both shear magnitude and direction. These unique, highly directional color-change responses of SSLCCs to aerodynamic shear allow for the full-surface visualization and measurement of continuous shear stress vector distributions. The objective of the present research was to investigate application of the SSLCC method through a transparent test surface. In this new back-light/back-view mode, the exposed surface of the SSLCC would be subjected to aerodynamic shear stress while the contact surface between the SSLCC and the solid, transparent wall would be illuminated and viewed in the same geometrical arrangement as applied in conventional applications. It was unknown at the outset whether or not color-change responses would be observable from the contact surface of the SSLCC, and, if seen, how these color-change responses might relate to those observed in standard practice.

New Techniques in Computational Aerodynamics.

DTIC Science & Technology

1987-08-06

the nature of a continuous range of nearby flow fields is of fundmental significance in the design and performance of aircraft. To treat this...parameters. The resulting matrix of derivatives of flow quantities is referred to as the Jacobi matrix. The subsequent procedure is in principle now...employed here also results in a method which is closely related to the method of characteristics. Special account must be taken of the appearance of

Sensitivity Analysis of the Static Aeroelastic Response of a Wing

NASA Technical Reports Server (NTRS)

Eldred, Lloyd B.

1993-01-01

A technique to obtain the sensitivity of the static aeroelastic response of a three dimensional wing model is designed and implemented. The formulation is quite general and accepts any aerodynamic and structural analysis capability. A program to combine the discipline level, or local, sensitivities into global sensitivity derivatives is developed. A variety of representations of the wing pressure field are developed and tested to determine the most accurate and efficient scheme for representing the field outside of the aerodynamic code. Chebyshev polynomials are used to globally fit the pressure field. This approach had some difficulties in representing local variations in the field, so a variety of local interpolation polynomial pressure representations are also implemented. These panel based representations use a constant pressure value, a bilinearly interpolated value. or a biquadraticallv interpolated value. The interpolation polynomial approaches do an excellent job of reducing the numerical problems of the global approach for comparable computational effort. Regardless of the pressure representation used. sensitivity and response results with excellent accuracy have been produced for large integrated quantities such as wing tip deflection and trim angle of attack. The sensitivities of such things as individual generalized displacements have been found with fair accuracy. In general, accuracy is found to be proportional to the relative size of the derivatives to the quantity itself.

Wind-Tunnel Investigation of Control-Surface Characteristics. 2 - A Large Aerodynamic Balance of Various Nose Shapes with a 30-Percent-Chord Flap on an NACA 0009 Airfoil

DTIC Science & Technology

1942-08-01

an extensive investigation of the aerodynamic characteristics of con- trol surfaces in order to provide data for desigD purposes...airfoil-flap combinations using plain flaps of various sises and with sealed gaps. (See references 2, 3, and 4.) The data ...made to provide sec- tion data for an airfoil having a flap with a large over- hang and to determine the effects of the

Influence matrix program for aerodynamic lifting surface theory. [in subsonic flows

NASA Technical Reports Server (NTRS)

Medan, R. T.; Ray, K. S.

1973-01-01

A users manual is described for a USA FORTRAN 4 computer program which computes an aerodynamic influence matrix and is one of several computer programs used to analyze lifting, thin wings in steady, subsonic flow according to a kernel function method lifting surface theory. The most significant features of the program are that it can treat unsymmetrical wings, control points can be placed on the leading and/or trailing edges, and a stable, efficient algorithm is used to compute the influence matrix.

Integrating aerodynamic surface modeling for computational fluid dynamics with computer aided structural analysis, design, and manufacturing

NASA Technical Reports Server (NTRS)

Thorp, Scott A.

1992-01-01

This presentation will discuss the development of a NASA Geometry Exchange Specification for transferring aerodynamic surface geometry between LeRC systems and grid generation software used for computational fluid dynamics research. The proposed specification is based on a subset of the Initial Graphics Exchange Specification (IGES). The presentation will include discussion of how the NASA-IGES standard will accommodate improved computer aided design inspection methods and reverse engineering techniques currently being developed. The presentation is in viewgraph format.

Aerodynamic repellency of impacting liquids

NASA Astrophysics Data System (ADS)

Gauthier, Anaïs; Bouillant, Ambre; Clanet, Christophe; Quéré, David

2018-05-01

Impacting liquids can be reflected by moving solid plates, provided the surface is fast enough. We describe and model here the threshold speed of bouncing, in particular as a function of the impact velocity of the incoming liquid. We also demonstrate that the aerodynamic force responsible for the nonwetting behavior induces an oblique rebound, which contributes to the liquid removal. In summary, this situation repels viscous, low surface tension drops of any size, all kinds of cases where repellency is impossible to achieve by other means.

Turbulence Control Through Selective Surface Heating Using Microwave Radiation

DTIC Science & Technology

2013-05-01

models. This type of plasma actuators needs further development to follow aerodynamic requirements of wind -tunnel experiments. 5. Ring -type plasma...modes of MW-heated elements in the aerodynamic experiment. Design of a resistive vibrator array for the airfoil model to be tested in a wind tunnel...

Aerodynamic properties of agricultural and natural surfaces in northwestern Tarim Basin

USDA-ARS?s Scientific Manuscript database

Friction velocity (u*) and aerodynamic roughness (z0) are important parameters that influence soil erosion, but no attempts have been made to quantify these parameters as affected by different land use types in the northwestern Tarim Basin. Wind velocity profiles were measured and used to determine ...

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.

Review of Cranked-Arrow Wing Aerodynamics Project: Its International Aeronautical Community Role

NASA Technical Reports Server (NTRS)

Lamar, John E.; Obara, Clifford J.

2007-01-01

This paper provides a brief history of the F-16XL-1 aircraft, its role in the High Speed Research (HSR) program and how it was morphed into the Cranked Arrow Wing Aerodynamics Project (CAWAP). Various flight, wind-tunnel and Computational Fluid Dynamics (CFD) data sets were generated during the CAWAP. These unique and open flight datasets for surface pressures, boundary-layer profiles and skinfriction distributions, along with surface flow data, are described and sample data comparisons given. This is followed by a description of how the project became internationalized to be known as Cranked Arrow Wing Aerodynamics Project International (CAWAPI) and is concluded by an introduction to the results of a 4 year CFD predictive study of data collected at flight conditions by participating researchers.

Sources of sound in fluid flows

NASA Technical Reports Server (NTRS)

Williams, J. E. F.

1974-01-01

Some features of a flow that produce acoustic radiation, particularly when the flow is turbulent and interacting with solid surfaces such as turbine or compressor blades are discussed. Early theoretical ideas on the subject are reviewed and are shown to be inadequate at high Mach number. Some recent theoretical developments that form the basis of a description of sound generation by supersonic flows interacting with surfaces are described. At high frequencies the problem is treated as one of describing the surface-induced diffraction field of adjacent aerodynamic quadrupole sources. This approach has given rise to distinctly new features of the problem that seem to have bearing on the radiating properties of relatively large aerodynamic surfaces.

Reconfigurable Control with Neural Network Augmentation for a Modified F-15 Aircraft

NASA Technical Reports Server (NTRS)

Burken, John J.; Williams-Hayes, Peggy; Kaneshige, John T.; Stachowiak, Susan J.

2006-01-01

Description of the performance of a simplified dynamic inversion controller with neural network augmentation follows. Simulation studies focus on the results with and without neural network adaptation through the use of an F-15 aircraft simulator that has been modified to include canards. Simulated control law performance with a surface failure, in addition to an aerodynamic failure, is presented. The aircraft, with adaptation, attempts to minimize the inertial cross-coupling effect of the failure (a control derivative anomaly associated with a jammed control surface). The dynamic inversion controller calculates necessary surface commands to achieve desired rates. The dynamic inversion controller uses approximate short period and roll axis dynamics. The yaw axis controller is a sideslip rate command system. Methods are described to reduce the cross-coupling effect and maintain adequate tracking errors for control surface failures. The aerodynamic failure destabilizes the pitching moment due to angle of attack. The results show that control of the aircraft with the neural networks is easier (more damped) than without the neural networks. Simulation results show neural network augmentation of the controller improves performance with aerodynamic and control surface failures in terms of tracking error and cross-coupling reduction.

Adaptive Control Using Neural Network Augmentation for a Modified F-15 Aircraft

NASA Technical Reports Server (NTRS)

Burken, John J.; Williams-Hayes, Peggy; Karneshige, J. T.; Stachowiak, Susan J.

2006-01-01

Description of the performance of a simplified dynamic inversion controller with neural network augmentation follows. Simulation studies focus on the results with and without neural network adaptation through the use of an F-15 aircraft simulator that has been modified to include canards. Simulated control law performance with a surface failure, in addition to an aerodynamic failure, is presented. The aircraft, with adaptation, attempts to minimize the inertial cross-coupling effect of the failure (a control derivative anomaly associated with a jammed control surface). The dynamic inversion controller calculates necessary surface commands to achieve desired rates. The dynamic inversion controller uses approximate short period and roll axis dynamics. The yaw axis controller is a sideslip rate command system. Methods are described to reduce the cross-coupling effect and maintain adequate tracking errors for control surface failures. The aerodynamic failure destabilizes the pitching moment due to angle of attack. The results show that control of the aircraft with the neural networks is easier (more damped) than without the neural networks. Simulation results show neural network augmentation of the controller improves performance with aerodynamic and control surface failures in terms of tracking error and cross-coupling reduction.

Influence of Additional Leading-Edge Surface Roughness on Performances in Highly Loaded Compressor Cascade

NASA Astrophysics Data System (ADS)

Chen, Shaowen; Xu, Hao; Sun, Shijun; Zhang, Longxin; Wang, Songtao

2015-05-01

Experimental research has been carried out at low speed to investigate the effect of additional leading-edge surface roughness on a highly-loaded axial compressor cascade. A 5-hole aerodynamic probe has been traversed across one pitch to obtain the distribution of total pressure loss coefficient, secondary flow vector, flow angles and other aerodynamic parameters at the exit section. Meanwhile, ink-trace flow visualization has been used to measure the flow fields on the walls of cascades and a detailed topology structure of the flow on the walls has been obtained. Aerodynamic parameters and flow characteristics are compared by arranging different levels of roughness on various parts of the leading edge. The results show that adding surface roughness at the leading edge and on the suction side obviously influences cascade performance. Aggravated 3-D flow separation significantly increases the loss in cascades, and the loss increases till 60% when the level of emery paper is 80 mm. Even there is the potential to improve cascade performance in local area of cascade passage. The influence of the length of surface roughness on cascade performance is not always adverse, and which depends on the position of surface roughness.

Forests tend to cool the land surface in the temperate zone: An analysis of the mechanisms controlling radiometric surface temperature change in managed temperate ecosystems

NASA Astrophysics Data System (ADS)

Stoy, P. C.; Katul, G. G.; Juang, J.; Siqueira, M. B.; Novick, K. A.; Essery, R.; Dore, S.; Kolb, T. E.; Montes-Helu, M. C.; Scott, R. L.

2010-12-01

Vegetation is an important control on the surface energy balance and thereby surface temperature. Boreal forests and arctic shrubs are thought to warm the land surface by absorbing more radiation than the vegetation they replace. The surface temperatures of tropical forests tend to be cooler than deforested landscapes due to enhanced evapotranspiration. The effects of reforestation on surface temperature change in the temperate zone is less-certain, but recent modeling efforts suggest forests have a global warming effect. We quantified the mechanisms driving radiometric surface changes following landcover changes using paired ecosystem case studies from the Ameriflux database with energy balance models of varying complexity. Results confirm previous findings that deciduous and coniferous forests in the southeastern U.S. are ca. 1 °C cooler than an adjacent field on an annual basis because aerodynamic/ecophysiological cooling of 2-3 °C outweighs an albedo-related warming of <1 °C. A 50-70% reduction in the aerodynamic resistance to sensible and latent heat exchange in the forests dominated the cooling effect. A grassland ecosystem that succeeded a stand-replacing ponderosa pine fire was ca. 1 °C warmer than unburned stands because a 1.5 °C aerodynamic warming offset a slight surface cooling due to greater albedo and soil heat flux. An ecosystem dominated by mesquite shrub encroachment was nearly 2 °C warmer than a native grassland ecosystem as aerodynamic and albedo-related warming outweighed a small cooling effect due to changes in soil heat flux. The forested ecosystems in these case studies are documented to have higher carbon uptake than the non-forested systems. Results suggest that temperate forests tend to cool the land surface and suggest that previous model-based findings that forests warm the Earth’s surface globally should be reconsidered.Changes to radiometric surface temperature (K) following changes in vegetation using paired ecosystem case studies C4 grassland and shrub ecosystem surface temperatures were adjusted for differences in air temperature across sites.

Aircraft Aerodynamic Parameter Detection Using Micro Hot-Film Flow Sensor Array and BP Neural Network Identification

PubMed Central

Que, Ruiyi; Zhu, Rong

2012-01-01

Air speed, angle of sideslip and angle of attack are fundamental aerodynamic parameters for controlling most aircraft. For small aircraft for which conventional detecting devices are too bulky and heavy to be utilized, a novel and practical methodology by which the aerodynamic parameters are inferred using a micro hot-film flow sensor array mounted on the surface of the wing is proposed. A back-propagation neural network is used to model the coupling relationship between readings of the sensor array and aerodynamic parameters. Two different sensor arrangements are tested in wind tunnel experiments and dependence of the system performance on the sensor arrangement is analyzed. PMID:23112638

Aircraft aerodynamic parameter detection using micro hot-film flow sensor array and BP neural network identification.

PubMed

Que, Ruiyi; Zhu, Rong

2012-01-01

Air speed, angle of sideslip and angle of attack are fundamental aerodynamic parameters for controlling most aircraft. For small aircraft for which conventional detecting devices are too bulky and heavy to be utilized, a novel and practical methodology by which the aerodynamic parameters are inferred using a micro hot-film flow sensor array mounted on the surface of the wing is proposed. A back-propagation neural network is used to model the coupling relationship between readings of the sensor array and aerodynamic parameters. Two different sensor arrangements are tested in wind tunnel experiments and dependence of the system performance on the sensor arrangement is analyzed.

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.

FLPP IXV Re-Entry Vehicle, Supersonic Charectisation Based on DNW SST Wind Tunnel Tests and CFD

NASA Astrophysics Data System (ADS)

Kapteijn, C.; Maseland, H.; Chiarelli, C.; Mareschi, V.; Tribot, J.-P.; Binetti, P.; Walloscheck, T.

2009-01-01

The European Space Agency ESA, has engaged in 2004, the IXV project (Intermediate eXperimental Vehicle) which is part of the FLPP (Future Launcher Preparatory Programme) aiming at answering to critical technological issues for controlled re-entry, while supporting the future generation launchers and to improve in general European capabilities in the strategic field of atmospheric re-entry for future space transportation, exploration and scientific applications. The IXV key mission and system objectives are the design, development, manufacturing, assembling and on- ground to in-flight verification of an autonomous European lifting and aerodynamically controlled re- entry system, integrating the critical re- entry technologies at the system level. In particular, the IXV shall demonstrate system integrated key technologies such as lifting flight control by means of aerodynamic surfaces that are one of the main primary objectives of the experimental investigation. Lifting and aerodynamic controlled re-entry represents a significant capability advancement with respect to the ballistic re-entry of capsules like the ARD. Since hypersonic aerodynamics is essentially different from supersonic aerodynamics, the current mission is to perform an atmospheric re-entry in combination with a safe recovery the in supersonic flight regime. However, mission extension to trimmed transonic flight is under consideration based on a preliminary analysis of the aerodynamic characteristics of the IXV configuration. Since the beginning of the IXV project, an aerodynamic data base (AEDB) has been built up and continuously updated integrating the additional information mainly provided by means of CFD (ie: Euler and Navier-Stokes) and lately also by means of WTTs. This AEDB serves for flying qualities analysis and for re-entry simulations. During the development phase B2/C1, the effectiveness of the control surfaces and their impact on te vehicle's aerodynamic forces in the supersonic regime is measured for a number of discrete deflection settings in the Super-Sonic wind Tunnel (SST) of DNW. Enabling an improved understanding of the measured aerodynamic characteristics, complementary computations were performed by Thales Alenia Space. The complete set of data was analyzed and compared enabling a consolidation of the nominal aerodynamic and aerodynamic uncertainties as well. The paper presents the main objectives of the supersonic characterisation of IXV including WTTs, and the main outcomes of the current data comparisons.

Aerodynamic characteristics of wings designed with a combined-theory method to cruise at a Mach number of 4.5

NASA Technical Reports Server (NTRS)

Mack, Robert J.

1988-01-01

A wind-tunnel study was conducted to determine the capability of a method combining linear theory and shock-expansion theory to design optimum camber surfaces for wings that will fly at high-supersonic/low-hypersonic speeds. Three force models (a flat-plate reference wing and two cambered and twisted wings) were used to obtain aerodynamic lift, drag, and pitching-moment data. A fourth pressure-orifice model was used to obtain surface-pressure data. All four wing models had the same planform, airfoil section, and centerbody area distribution. The design Mach number was 4.5, but data were also obtained at Mach numbers of 3.5 and 4.0. Results of these tests indicated that the use of airfoil thickness as a theoretical optimum, camber-surface design constraint did not improve the aerodynamic efficiency or performance of a wing as compared with a wing that was designed with a zero-thickness airfoil (linear-theory) constraint.

Results of tests OA12 and IA9 in the Ames Research Center unitary plan wind tunnels on an 0.030-scale model of the space shuttle vehicle 2A to determine aerodynamic loads, volume 14

NASA Technical Reports Server (NTRS)

Spangler, R. H.

1974-01-01

Tests were conducted in wind tunnels during April and May 1973, on a 0.030-scale replica of the Space Shuttle Vehicle Configuration 2A. Aerodynamic loads data were obtained at Mach numbers from 0.6 to 3.5. The investigation included tests on the integrated (launch) configuration and the isolated orbiter (entry configuration). The integrated vehicle was tested at angles of attack and sideslip from -8 degrees to +8 degrees. The isolated orbiter was tested at angles of attack from -15 degrees to +40 degrees and angles of sideslip from -10 degrees to +10 degrees as dictated by trajectory considerations. The effects of orbiter/external tank incidence angle and deflected control surfaces on aerodynamic loads were also investigated. Tabulated pressure data were obtained for upper and lower wing surfaces and left and right vertical tail surfaces.

A study of flow past an airfoil with a jet issuing from its lower surface

NASA Technical Reports Server (NTRS)

Krothapalli, A.; Leopold, D.

1984-01-01

The aerodynamics of a NACA 0018 airfoil with a rectangular jet of finite aspect ratio exiting from its lower surface at 90 deg to the chord were investigated. The jet was located at 50% of the wing chord. Measurements include static pressures on the airfoil surface, total pressures in the near wake, and local velocity vectors in different planes of the wake. The effects of jet cross flow interaction on the aerodynamics of the airfoil are studied. It is indicated that at all values of momentum coefficients, the jet cross flow interaction produces a strong contra-rotating vortex structure in the near wake. The flow behind the jet forms a closed recirculation region which extends up to a chord length down stream of the trailing edge which results in the flow field to become highly three dimensional. The various aerodynamic force coefficients vary significantly along the span of the wing. The results are compared with a jet flap configuration.

Aerodynamic Parameter Estimation for the X-43A (Hyper-X) from Flight Data

NASA Technical Reports Server (NTRS)

Morelli, Eugene A.; Derry, Stephen D.; Smith, Mark S.

2005-01-01

Aerodynamic parameters were estimated based on flight data from the third flight of the X-43A hypersonic research vehicle, also called Hyper-X. Maneuvers were flown using multiple orthogonal phase-optimized sweep inputs applied as simultaneous control surface perturbations at Mach 8, 7, 6, 5, 4, and 3 during the vehicle descent. Aerodynamic parameters, consisting of non-dimensional longitudinal and lateral stability and control derivatives, were estimated from flight data at each Mach number. Multi-step inputs at nearly the same flight conditions were also flown to assess the prediction capability of the identified models. Prediction errors were found to be comparable in magnitude to the modeling errors, which indicates accurate modeling. Aerodynamic parameter estimates were plotted as a function of Mach number, and compared with estimates from the pre-flight aerodynamic database, which was based on wind-tunnel tests and computational fluid dynamics. Agreement between flight estimates and values computed from the aerodynamic database was excellent overall.

Rotationally Adaptive Flight Test Surface

NASA Technical Reports Server (NTRS)

Barrett, Ron

1999-01-01

Research on a new design of flutter exciter vane using adaptive materials was conducted. This novel design is based on all-moving aerodynamic surface technology and consists of a structurally stiff main spar, a series of piezoelectric actuator elements and an aerodynamic shell which is pivoted around the main spar. The work was built upon the current missile-type all-moving surface designs and change them so they are better suited for flutter excitation through the transonic flight regime. The first portion of research will be centered on aerodynamic and structural modeling of the system. USAF DatCom and vortex lattice codes was used to capture the fundamental aerodynamics of the vane. Finite element codes and laminated plate theory and virtual work analyses will be used to structurally model the aerodynamic vane and wing tip. Following the basic modeling, a flutter test vane was designed. Each component within the structure was designed to meet the design loads. After the design loads are met, then the deflections will be maximized and the internal structure will be laid out. In addition to the structure, a basic electrical control network will be designed which will be capable of driving a scaled exciter vane. The third and final stage of main investigation involved the fabrication of a 1/4 scale vane. This scaled vane was used to verify kinematics and structural mechanics theories on all-moving actuation. Following assembly, a series of bench tests was conducted to determine frequency response, electrical characteristics, mechanical and kinematic properties. Test results indicate peak-to-peak deflections of 1.1 deg with a corner frequency of just over 130 Hz.

Owl-inspired leading-edge serrations play a crucial role in aerodynamic force production and sound suppression.

PubMed

Rao, Chen; Ikeda, Teruaki; Nakata, Toshiyuki; Liu, Hao

2017-07-04

Owls are widely known for silent flight, achieving remarkably low noise gliding and flapping flights owing to their unique wing morphologies, which are normally characterized by leading-edge serrations, trailing-edge fringes and velvet-like surfaces. How these morphological features affect aerodynamic force production and sound suppression or noise reduction, however, is still not well known. Here we address an integrated study of owl-inspired single feather wing models with and without leading-edge serrations by combining large-eddy simulations (LES) with particle-image velocimetry (PIV) and force measurements in a low-speed wind tunnel. With velocity and pressure spectra analysis, we demonstrate that leading-edge serrations can passively control the laminar-turbulent transition over the upper wing surface, i.e. the suction surface at all angles of attack (0°  <  AoA  <  20°), and hence play a crucial role in aerodynamic force and sound production. We find that there exists a tradeoff between force production and sound suppression: serrated leading-edges reduce aerodynamic performance at lower AoAs  <  15° compared to clean leading-edges but are capable of achieving both noise reduction and aerodynamic performance at higher AoAs  >  15° where owl wings often reach in flight. Our results indicate that the owl-inspired leading-edge serrations may be a useful device for aero-acoustic control in biomimetic rotor designs for wind turbines, aircrafts, multi-rotor drones as well as other fluid machinery.

The Effects of Streamwise-Deflected Wing Tips on the Aerodynamic Characteristics of an Aspect Ratio-2 Triangular Wing, Body, and Tail Combination

NASA Technical Reports Server (NTRS)

Peterson, Victor L.

1959-01-01

An investigation has been conducted on a triangular wing and body combination to determine the effects on the aerodynamic characteristics resulting from deflecting portions of the wing near the tips 900 to the wing surface about streamwise hinge lines. Experimental data were obtained for Mach numbers of 0.70, 1.30, 1.70, and 2.22 and for angles of attack ranging from -5 deg to +18 deg at sideslip angles of 0 deg and 5 deg. The results showed that the aerodynamic center shift experienced by the triangular wing and body combination as the Mach number was increased from subsonic to supersonic could be reduced by about 40 percent by deflecting the outboard 4 percent of the total area of each wing panel. Deflection about the same hinge line of additional inboard surfaces consisting of 2 percent of the total area of each wing panel resulted in a further reduction of the aerodynamic center travel of 10 percent. The resulting reductions in the stability were accompanied by increases in the drag due to lift and, for the case of the configuration with all surfaces deflected, in the minimum drag. The combined effects of reduced stability and increased drag of the untrimmed configuration on the trimmed lift-drag ratios were estimated from an analysis of the cases in which the wing-body combination with or without tips deflected was assumed to be controlled by a canard. The configurations with deflected surfaces had higher trimmed lift-drag ratios than the model with undeflected surfaces at Mach numbers up to about 1.70. Deflecting either the outboard surfaces or all of the surfaces caused the directional stability to be increased by increments that were approximately constant with increasing angle of attack at each Mach number. The effective dihedral was decreased at all angles of attack and Mach numbers when the surfaces were deflected.

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.

Development of subminiature multi-sensor hot-wire probes

NASA Technical Reports Server (NTRS)

Westphal, Russell V.; Ligrani, Phillip M.; Lemos, Fred R.

1988-01-01

Limitations on the spatial resolution of multisensor hot wire probes have precluded accurate measurements of Reynolds stresses very near solid surfaces in wind tunnels and in many practical aerodynamic flows. The fabrication, calibration and qualification testing of very small single horizontal and X-array hot-wire probes which are intended to be used near solid boundaries in turbulent flows where length scales are particularly small, is described. Details of the sensor fabrication procedure are reported, along with information needed to successfully operate the probes. As compared with conventional probes, manufacture of the subminiature probes is more complex, requiring special equipment and careful handling. The subminiature probes tested were more fragile and shorter lived than conventional probes; they obeyed the same calibration laws but with slightly larger experimental uncertainty. In spite of these disadvantages, measurements of mean statistical quantities and spectra demonstrate the ability of the subminiature sensors to provide the measurements in the near wall region of turbulent boundary layers that are more accurate than conventional sized probes.

TRW vortex-lattice method subsonic aerodynamic analysis for multiple-lifting-surfaces (N. surface) TRW program number HA010B

NASA Technical Reports Server (NTRS)

Gomez, A. V.

1972-01-01

The program was designed to provide solutions of engineering accuracy for determining the aerodynamic loads on single- or multiple-lifting-surface configurations that represent vehicles in subsonic flight, e.g., wings, wing-tail, wing-canard, lifting bodies, etc. The preparation is described of the input data, associated input arrangement, and the output format for the program data, including specification of the various operational details of the program such as array sizes, tape numbers utilized, and program dumps. A full description of the underlying theory used in the program development and a review of the program qualification tests are included.

On problems of analyzing aerodynamic properties of blunted rotary bodies with small random surface distortions under supersonic and hypersonic flows

NASA Astrophysics Data System (ADS)

Degtyar, V. G.; Kalashnikov, S. T.; Mokin, Yu. A.

2017-10-01

The paper considers problems of analyzing aerodynamic properties (ADP) of reenetry vehicles (RV) as blunted rotary bodies with small random surface distortions. The interactions of math simulation of surface distortions, selection of tools for predicting ADPs of shaped bodies, evaluation of different-type ADP variations and their adaptation for dynamic problems are analyzed. The possibilities of deterministic and probabilistic approaches to evaluation of ADP variations are considered. The practical value of the probabilistic approach is demonstrated. The examples of extremal deterministic evaluations of ADP variations for a sphere and a sharp cone are given.

Computer Code Aids Design Of Wings

NASA Technical Reports Server (NTRS)

Carlson, Harry W.; Darden, Christine M.

1993-01-01

AERO2S computer code developed to aid design engineers in selection and evaluation of aerodynamically efficient wing/canard and wing/horizontal-tail configurations that includes simple hinged-flap systems. Code rapidly estimates longitudinal aerodynamic characteristics of conceptual airplane lifting-surface arrangements. Developed in FORTRAN V on CDC 6000 computer system, and ported to MS-DOS environment.

Aerodynamic Measurements on a Large Splitter Plate for the NASA Langley Transonic Dynamics Tunnel

NASA Technical Reports Server (NTRS)

Schuster, David M.

2001-01-01

Tests conducted in the NASA Langley Research Center Transonic Dynamics Tunnel (TDT) assess the aerodynamic characteristics of a splitter plate used to test some semispan models in this facility. Aerodynamic data are analyzed to determine the effect of the splitter plate on the operating characteristics of the TDT, as well as to define the range of conditions over which the plate can be reasonably used to obtain aerodynamic data. Static pressures measurements on the splitter plate surface and the equipment fairing between the wind tunnel wall and the splitter plate are evaluated to determine the flow quality around the apparatus over a range of operating conditions. Boundary layer rake data acquired near the plate surface define the viscous characteristics of the flow over the plate. Data were acquired over a range of subsonic, transonic and supersonic conditions at dynamic pressures typical for models tested on this apparatus. Data from this investigation should be used as a guide for the design of TDT models and tests using the splitter plate, as well as to guide future splitter plate design for this facility.

Static aeroelastic deformation of flexible skin for continuous variable trailing-edge camber wing

NASA Astrophysics Data System (ADS)

Liu, Libo; Yin, Weilong; Dai, Fuhong; Liu, Yanju; Leng, Jinsong

2011-03-01

The method for analyzing the static aeroelastic deformation of flexible skin under the air loads was developed. The effect of static aeroelastic deformation of flexible skin on the aerodynamic characteristics of aerofoil and the design parameters of skin was discussed. Numerical results show that the flexible skin on the upper surface of trailing-edge will bubble under the air loads and the bubble has a powerful effect on the aerodynamic pressure near the surface of local deformation. The static aeroelastic deformation of flexible skin significantly affects the aerodynamic characteristics of aerofoil. At small angle of attack, the drag coefficient increases and the lift coefficient decreases. With the increasing angle of attack, the effect of flexible skin on the aerodynamic characteristics of aerofoil is smaller and smaller. The deformation of flexible skin becomes larger and larger with the free-stream velocity increasing. When the free-stream velocity is greater than a value, both of the deformation of flexible skin and the drag coefficient of aerofoil increase rapidly. The maximum tensile strain of flexible skin is increased with consideration of the static aeroelastic deformation.

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.

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.

Incompressible lifting-surface aerodynamics for a rotor-stator combination

NASA Technical Reports Server (NTRS)

Ramachandra, S. M.

1984-01-01

Current literature on the three dimensional flow through compressor cascades deals with a row of rotor blades in isolation. Since the distance between the rotor and stator is usually 10 to 20 percent of the blade chord, the aerodynamic interference between them has to be considered for a proper evaluation of the aerothermodynamic performance of the stage. A unified approach to the aerodynamics of the incompressible flow through a stage is presented that uses the lifting surface theory for a compressor cascade of arbitrary camber and thickness distribution. The effects of rotor stator interference are represented as a linear function of the rotor and stator flows separately. The loading distribution on the rotor and stator flows separately. The loading distribution on the rotor and stator blades and the interference factor are determined concurrently through a matrix iteration process.

Aerodynamic heating and surface temperatures on vehicles for computer-aided design studies

NASA Technical Reports Server (NTRS)

Dejarnette, F. R.; Kania, L. A.; Chitty, A.

1983-01-01

A computer subprogram has been developed to calculate aerodynamic and radiative heating rates and to determine surface temperatures by integrating the heating rates along the trajectory of a vehicle. Convective heating rates are calculated by applying the axisymmetric analogue to inviscid surface streamlines and using relatively simple techniques to calculate laminar, transitional, or turbulent heating rates. Options are provided for the selection of gas model, transition criterion, turbulent heating method, Reynolds Analogy factor, and entropy-layer swallowing effects. Heating rates are compared to experimental data, and the time history of surface temperatures are given for a high-speed trajectory. The computer subprogram is developed for preliminary design and mission analysis where parametric studies are needed at all speeds.

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.

Computation of flows in a turn-around duct and a turbine cascade using advanced turbulence models

NASA Technical Reports Server (NTRS)

Lakshminarayana, B.; Luo, J.

1993-01-01

Numerical investigation has been carried out to evaluate the capability of the Algebraic Reynolds Stress Model (ARSM) and the Nonlinear Stress Model (NLSM) to predict strongly curved turbulent flow in a turn-around duct (TAD). The ARSM includes the near-wall damping term of pressure-strain correlation phi(sub ij,w), which enables accurate prediction of individual Reynolds stress components in wall flows. The TAD mean flow quantities are reasonably well predicted by various turbulence models. The ARSM yields better predictions for both the mean flow and the turbulence quantities than the NLSM and the k-epsilon (k = turbulent kinetic energy, epsilon = dissipation rate of k) model. The NLSM also shows slight improvement over the k-epsilon model. However, all the models fail to capture the recovery of the flow from strong curvature effects. The formulation for phi(sub ij,w) appears to be incorrect near the concave surface. The hybrid k-epsilon/ARSM, Chien's k-epsilon model, and Coakley's q-omega (q = the square root of k, omega = epsilon/k) model have also been employed to compute the aerodynamics and heat transfer of a transonic turbine cascade. The surface pressure distributions and the wake profiles are predicted well by all the models. The k-epsilon model and the k-epsilon/ARSM model provide better predictions of heat transfer than the q-omega model. The k-epsilon/ARSM solutions show significant differences in the predicted skin friction coefficients, heat transfer rates and the cascade performance parameters, as compared to the k-epsilon model. The k-epsilon/ARSM model appears to capture, qualitatively, the anisotropy associated with by-pass transition.

A two-dimensional cascade solution using minimized surface singularity density distributions - with application to film cooled turbine blades

NASA Technical Reports Server (NTRS)

Mcfarland, E.; Tabakoff, W.; Hamed, A.

1977-01-01

An investigation of the effects of coolant injection on the aerodynamic performance of cooled turbine blades is presented. The coolant injection is modeled in the inviscid irrotational adiabatic flow analysis through the cascade using the distributed singularities approach. The resulting integral equations are solved using a minimized surface singularity density criteria. The aerodynamic performance was evaluated using this solution in conjunction with an existing mixing theory analysis. The results of the present analysis are compared with experimental measurements in cold flow tests.

Experimental studies on the aerodynamic performance and dynamic response of flow direction sensing vanes

NASA Technical Reports Server (NTRS)

Barna, P. S.; Crossman, G. R.

1976-01-01

Systematic investigations were performed on a variety of aerodynamic surfaces to obtain their potential for possible application to wind vanes. Among the surfaces tested were: (1) single vanes consisting of flat plates of various planforms having aspect ratios between 0.5 and 5; (2) bi-vanes with aspect ratio 2.5; (3) various cone and box vanes; and (4) various cruciform configurations. The models were subjected to windtunnel tests. In addition to lift and drag force measurements, damping and frequency test were performed under a variety of flow conditions.

Photogrammetry of a Hypersonic Inflatable Aerodynamic Decelerator

NASA Technical Reports Server (NTRS)

Kushner, Laura Kathryn; Littell, Justin D.; Cassell, Alan M.

2013-01-01

In 2012, two large-scale models of a Hypersonic Inflatable Aerodynamic decelerator were tested in the National Full-Scale Aerodynamic Complex at NASA Ames Research Center. One of the objectives of this test was to measure model deflections under aerodynamic loading that approximated expected flight conditions. The measurements were acquired using stereo photogrammetry. Four pairs of stereo cameras were mounted inside the NFAC test section, each imaging a particular section of the HIAD. The views were then stitched together post-test to create a surface deformation profile. The data from the photogram- metry system will largely be used for comparisons to and refinement of Fluid Structure Interaction models. This paper describes how a commercial photogrammetry system was adapted to make the measurements and presents some preliminary results.

Aerodynamic heating in transitional hypersonic boundary layers: Role of second-mode instability

NASA Astrophysics Data System (ADS)

Zhu, Yiding; Chen, Xi; Wu, Jiezhi; Chen, Shiyi; Lee, Cunbiao; Gad-el-Hak, Mohamed

2018-01-01

The evolution of second-mode instabilities in hypersonic boundary layers and its effects on aerodynamic heating are investigated. Experiments are conducted in a Mach 6 wind tunnel using fast-response pressure sensors, fluorescent temperature-sensitive paint, and particle image velocimetry. Calculations based on parabolic stability equations and direct numerical simulations are also performed. It is found that second-mode waves, accompanied by high-frequency alternating fluid compression and expansion, produce intense aerodynamic heating in a small region that rapidly heats the fluid passing through it. As the second-mode waves decay downstream, the dilatation-induced aerodynamic heating decreases while its shear-induced counterpart keeps growing. The latter brings about a second growth of the surface temperature when transition is completed.

Effects of several factors on theoretical predictions of airplane spin characteristics. [dynamic models

NASA Technical Reports Server (NTRS)

Bihrle, W., Jr.; Barnhart, B.

1974-01-01

The influence of different mathematical and aerodynamic models on computed spin motion was investigated along with the importance of some of the aerodynamic and nonaerodynamic quantities defined in these models. An analytical technique was used which included the aerodynamic forces and moments acting on a spinning aircraft due to steady rotational flow and the contribution of the rotary derivatives to the oscillatory component of the total angular rates. It was shown that (1) during experimental-analytical correlation studies, the flight-recorded control time histories must be faithfully duplicated since the spinning motion can be sensitive to a small change in the application of the spin entry controls; (2) an error in the assumed inertias, yawing moments at high angle of attack, and initial spin entry bank angle do not influence the developed spin significantly; (3) damping in pitch derivatives and the center of gravity location play a role in the spinning motion; and (4) the experimental spin investigations conducted in a constant atmospheric density environment duplicate the Froude number only at the initial full-scale spin altitude (since the full-scale airplane at high altitudes experiences large density changes during the spin.)

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.

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.

Supersonic Flight Dynamics Test: Trajectory, Atmosphere, and Aerodynamics Reconstruction

NASA Technical Reports Server (NTRS)

Kutty, Prasad; Karlgaard, Christopher D.; Blood, Eric M.; O'Farrell, Clara; Ginn, Jason M.; Shoenenberger, Mark; Dutta, Soumyo

2015-01-01

The Supersonic Flight Dynamics Test is a full-scale flight test of a Supersonic Inflatable Aerodynamic Decelerator, which is part of the Low Density Supersonic Decelerator technology development project. The purpose of the project is to develop and mature aerodynamic decelerator technologies for landing large mass payloads on the surface of Mars. The technologies include a Supersonic Inflatable Aerodynamic Decelerator and Supersonic Parachutes. The first Supersonic Flight Dynamics Test occurred on June 28th, 2014 at the Pacific Missile Range Facility. This test was used to validate the test architecture for future missions. The flight was a success and, in addition, was able to acquire data on the aerodynamic performance of the supersonic inflatable decelerator. This paper describes the instrumentation, analysis techniques, and acquired flight test data utilized to reconstruct the vehicle trajectory, atmosphere, and aerodynamics. The results of the reconstruction show significantly higher lofting of the trajectory, which can partially be explained by off-nominal booster motor performance. The reconstructed vehicle force and moment coefficients fall well within pre-flight predictions. A parameter identification analysis indicates that the vehicle displayed greater aerodynamic static stability than seen in pre-flight computational predictions and ballistic range tests.

Temporal dynamics of salt crust patterns on a sodic playa: implications for aerodynamic roughness and dust emission potential

NASA Astrophysics Data System (ADS)

Nield, Joanna; Bryant, Robert; Wiggs, Giles; King, James; Thomas, David; Eckardt, Frank; Washington, Richard

2015-04-01

Salt pans (or playas) are common in arid environments and can be major sources of windblown mineral dust, but there are uncertainties associated with their dust emission potential. These landforms typically form crusts which modify both their erosivity and erodibility by limiting sediment availability, modifying surface and aerodynamic roughness and limiting evaporation rates and sediment production. Here we show the relationship between seasonal surface moisture change and crust pattern development on part of the Makgadikgadi Pans of Botswana (a Southern Hemisphere playa that emits significant dust), based on both remote-sensing and field surface and atmospheric measurements. We use high resolution (sub-cm) terrestrial laser scanning (TLS) surveys over weekly, monthly and annual timescales to accurately characterise crustal ridge thrusting and collapse. Ridge development can change surface topography as much as 30 mm/week on fresh pan areas that have recently been reset by flooding. The corresponding change aerodynamic roughness can be as much as 3 mm/week. At the same time, crack densities across the surface increase and this raises the availability of erodible fluffy, low density dust source sediment stored below the crust layer. We present a conceptual model accounting for the driving forces (subsurface, surface and atmospheric moisture) and feedbacks between these and surface shape that lead to crust pattern trajectories between highly emissive degraded surfaces and less emissive ridged or continuous crusts. These findings improve our understanding of temporal changes in dust availability and supply from playa source regions.

Adhesion and aerodynamic forces for the resuspension of non-spherical particles in outdoor environments

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

Brambilla, Sara; Speckart, Scott; Brown, Michael J.

Particles deposited on an outdoor surface can be resuspended by wind gusts, become airborne, and be inhaled if small enough. If toxic or infectious, these particles may be dangerous for the populace health. It is therefore important to determine under which weather conditions a deposit of particle could be resuspended to implement the best response actions and plan clean-up. To this scope, one needs to consider the competing forces acting on the particle keeping it attached to the surface (gravity and adhesion) or trying to remove it (aerodynamic forces, i.e., lift and drag). Here, this article reviews the current understandingmore » of the aforementioned forces for colloidal spherical particles and extends the existing theories to rod-shaped particles, representative for instance of Bacillus spores. In particular, for the adhesion force, the Derjaguin approximation was used and the adhesion force was computed from the radii of curvature of the particle and the surface at the point of closest approach. For the aerodynamic forces, we re-derived the equations for the drag and lift forces accounting for the shape of the particle. Both smooth and rough surfaces will be discussed, the former as idealized cases, the latter as more representative of real outdoor surfaces.« less

Adhesion and aerodynamic forces for the resuspension of non-spherical particles in outdoor environments

DOE PAGES

Brambilla, Sara; Speckart, Scott; Brown, Michael J.

2017-07-26

Particles deposited on an outdoor surface can be resuspended by wind gusts, become airborne, and be inhaled if small enough. If toxic or infectious, these particles may be dangerous for the populace health. It is therefore important to determine under which weather conditions a deposit of particle could be resuspended to implement the best response actions and plan clean-up. To this scope, one needs to consider the competing forces acting on the particle keeping it attached to the surface (gravity and adhesion) or trying to remove it (aerodynamic forces, i.e., lift and drag). Here, this article reviews the current understandingmore » of the aforementioned forces for colloidal spherical particles and extends the existing theories to rod-shaped particles, representative for instance of Bacillus spores. In particular, for the adhesion force, the Derjaguin approximation was used and the adhesion force was computed from the radii of curvature of the particle and the surface at the point of closest approach. For the aerodynamic forces, we re-derived the equations for the drag and lift forces accounting for the shape of the particle. Both smooth and rough surfaces will be discussed, the former as idealized cases, the latter as more representative of real outdoor surfaces.« less

Improvement in Capsule Abort Performance Using Supersonic Aerodynamic Interaction by Fences

NASA Astrophysics Data System (ADS)

Koyama, Hiroto; Wang, Yunpeng; Ozawa, Hiroshi; Doi, Katsunori; Nakamura, Yoshiaki

The space transportation system will need advanced abort systems to secure crew against serious accidents. Here this study deals with the capsule-type space transportation systems with a Launch Abort System (LAS). This system is composed of a conic capsule as a Launch Abort Vehicle (LAV) and a cylindrical rocket as a Service Module (SM), and the capsule is moved away from the rocket by supersonic aerodynamic interactions in an emergency. We propose a method to improve the performance of the LAV by installing fences at the edges of surfaces on the rocket and capsule sides. Their effects were investigated by experimental measurements and numerical simulations. Experimental results show that the fences on the rocket and capsule surfaces increase the aerodynamic thrust force on the capsule by 70% in a certain clearance between the capsule and rocket. Computational results show the detailed flow fields where the centripetal flow near the surface on the rocket side is induced by the fence on the rocket side and the centrifugal flow near the surface on the capsule side is blocked by the fence on the capsule side. These results can confirm favorable effects of the fences on the performance of the LAS.

Fixture For Mounting A Pressure Sensor

NASA Technical Reports Server (NTRS)

Cagle, Christopher M.

1995-01-01

Fixture for mounting pressure sensor in aerodynamic model simplifies task of removal and replacement of sensor in event sensor becomes damaged. Makes it unnecessary to dismantle model. Also minimizes any change in aerodynamic characteristics of model in event of replacement. Removable pressure sensor installed in fixture in wall of model. Wires from sensor pass through channel under surface.

Determining Surface Roughness in Urban Areas Using Lidar Data

NASA Technical Reports Server (NTRS)

Holland, Donald

2009-01-01

An automated procedure has been developed to derive relevant factors, which can increase the ability to produce objective, repeatable methods for determining aerodynamic surface roughness. Aerodynamic surface roughness is used for many applications, like atmospheric dispersive models and wind-damage models. For this technique, existing lidar data was used that was originally collected for terrain analysis, and demonstrated that surface roughness values can be automatically derived, and then subsequently utilized in disaster-management and homeland security models. The developed lidar-processing algorithm effectively distinguishes buildings from trees and characterizes their size, density, orientation, and spacing (see figure); all of these variables are parameters that are required to calculate the estimated surface roughness for a specified area. By using this algorithm, aerodynamic surface roughness values in urban areas can then be extracted automatically. The user can also adjust the algorithm for local conditions and lidar characteristics, like summer/winter vegetation and dense/sparse lidar point spacing. Additionally, the user can also survey variations in surface roughness that occurs due to wind direction; for example, during a hurricane, when wind direction can change dramatically, this variable can be extremely significant. In its current state, the algorithm calculates an estimated surface roughness for a square kilometer area; techniques using the lidar data to calculate the surface roughness for a point, whereby only roughness elements that are upstream from the point of interest are used and the wind direction is a vital concern, are being investigated. This technological advancement will improve the reliability and accuracy of models that use and incorporate surface roughness.

Results of the Imager for Mars Pathfinder windsock experiment

USGS Publications Warehouse

Sullivan, R.; Greeley, R.; Kraft, M.; Wilson, G.; Golombek, M.; Herkenhoff, K.; Murphy, J.; Smith, P.

2000-01-01

The Imager for Mars Pathfinder (IMP) windsock experiment measured wind speeds at three heights within 1.2 m of the Martian surface during Pathfinder landed operations. These wind data allowed direct measurement of near-surface wind profiles on Mars for the first time, including determination of aerodynamic roughness length and wind friction speeds. Winds were light during periods of windsock imaging, but data from the strongest breezes indicate aerodynamic roughness length of 3 cm at the landing site, with wind friction speeds reaching 1 m/s. Maximum wind friction speeds were about half of the threshold-of-motion friction speeds predicted for loose, fine-grained materials on smooth Martian terrain and about one third of the threshold-of-motion friction speeds predicted for the same size particles over terrain with aerodynamic roughness of 3 cm. Consistent with this, and suggesting that low wind speeds prevailed when the windsock array was not imaged and/or no particles were available for aeolian transport, no wind-related changes to the surface during mission operations have been recognized. The aerodynamic roughness length reported here implies that proposed deflation of fine particles around the landing site, or activation of duneforms seen by IMP and Sojourner, would require wind speeds >28 m/s at the Pathfinder top windsock height (or >31 m/s at the equivalent Viking wind sensor height of 1.6 m) and wind speeds >45 m/s above 10 m. These wind speeds would cause rock abrasion if a supply of durable particles were available for saltation. Previous analyses indicate that the Pathfinder landing site probably is rockier and rougher than many other plains units on Mars, so aerodynamic roughness length elsewhere probably is less than the 3-cm value reported for the Pathfinder site. Copyright 2000 by the American Geophysical Union.

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

Experimental Aerodynamic Characteristics of a Joined-wing Research Aircraft Configuration

NASA Technical Reports Server (NTRS)

Smith, Stephen C.; Stonum, Ronald K.

1989-01-01

A wind-tunnel test was conducted at Ames Research Center to measure the aerodynamic characteristics of a joined-wing research aircraft (JWRA). This aircraft was designed to utilize the fuselage and engines of the existing NASA AD-1 aircraft. The JWRA was designed to have removable outer wing panels to represent three different configurations with the interwing joint at different fractions of the wing span. A one-sixth-scale wind-tunnel model of all three configurations of the JWRA was tested in the Ames 12-Foot Pressure Wind Tunnel to measure aerodynamic performance, stability, and control characteristics. The results of these tests are presented. Longitudinal and lateral-directional characteristics were measured over an angle of attack range of -7 to 14 deg and over an angle of sideslip range of -5 to +2.5 deg at a Mach number of 0.35 and a Reynolds number of 2.2x10(6)/ft. Various combinations of deflected control surfaces were tested to measure the effectiveness and impact on stability of several control surface arrangements. In addition, the effects on stall and post-stall aerodynamic characteristics from small leading-edge devices called vortilons were measured. The results of these tests indicate that the JWRA had very good aerodynamic performance and acceptable stability and control throughout its flight envelope. The vortilons produced a profound improvement in the stall and post-stall characteristics with no measurable effects on cruise performance.

CFD Simulations in Support of Shuttle Orbiter Contingency Abort Aerodynamic Database Enhancement

NASA Technical Reports Server (NTRS)

Papadopoulos, Periklis E.; Prabhu, Dinesh; Wright, Michael; Davies, Carol; McDaniel, Ryan; Venkatapathy, E.; Wercinski, Paul; Gomez, R. J.

2001-01-01

Modern Computational Fluid Dynamics (CFD) techniques were used to compute aerodynamic forces and moments of the Space Shuttle Orbiter in specific portions of contingency abort trajectory space. The trajectory space covers a Mach number range of 3.5-15, an angle-of-attack range of 20deg-60deg, an altitude range of 100-190 kft, and several different settings of the control surfaces (elevons, body flap, and speed brake). Presented here are details of the methodology and comparisons of computed aerodynamic coefficients against the values in the current Orbiter Operational Aerodynamic Data Book (OADB). While approximately 40 cases have been computed, only a sampling of the results is provided here. The computed results, in general, are in good agreement with the OADB data (i.e., within the uncertainty bands) for almost all the cases. However, in a limited number of high angle-of-attack cases (at Mach 15), there are significant differences between the computed results, especially the vehicle pitching moment, and the OADB data. A preliminary analysis of the data from the CFD simulations at Mach 15 shows that these differences can be attributed to real-gas/Mach number effects. The aerodynamic coefficients and detailed surface pressure distributions of the present simulations are being used by the Shuttle Program in the evaluation of the capabilities of the Orbiter in contingency abort scenarios.

Advanced Response Surface Modeling of Ares I Roll Control Jet Aerodynamic Interactions

NASA Technical Reports Server (NTRS)

Favaregh, Noah M.

2010-01-01

The Ares I rocket uses roll control jets. These jets have aerodynamic implications as they impinge on the surface and protuberances of the vehicle. The jet interaction on the body can cause an amplification or a reduction of the rolling moment produced by the jet itself, either increasing the jet effectiveness or creating an adverse effect. A design of experiments test was planned and carried out using computation fluid dynamics, and a subsequent response surface analysis ensued on the available data to characterize the jet interaction across the ascent portion of the Ares I flight envelope. Four response surface schemes were compared including a single response surface covering the entire design space, separate sector responses that did not overlap, continuously overlapping surfaces, and recursive weighted response surfaces. These surfaces were evaluated on traditional statistical metrics as well as visual inspection. Validation of the recursive weighted response surface was performed using additionally available data at off-design point locations.

Surface thermochemical effects on TPS-coupled aerothermodynamics in hypersonic Martian gas flow

NASA Astrophysics Data System (ADS)

Yang, Xiaofeng; Gui, Yewei; Tang, Wei; Du, Yanxia; Liu, Lei; Xiao, Guangming; Wei, Dong

2018-06-01

This paper deals with the surface thermochemical effects on TPS-coupled aerothermodynamics in hypersonic Martian gas flow. An interface condition with finite-rate thermochemistry was established to balance the three-dimensional Navier-Stokes solver and TPS thermal response solver, and a series of coupled simulations of chemical non-equilibrium aerothermodynamics and structure heat transfer with various surface catalycities were performed for hypersonic Mars entries. The analysis of surface thermochemistry reveals that the surface chemical reactions have great contribution to aerodynamic heating, and the temperature-dependence of finite-rate catalysis highly influences the evolution of the coupling aerodynamic heating in the coupling process. For fixed free stream parameters with proper catalytic excitation energy, a "leap" phenomenon of the TPS-coupled heat flux with the coupling time appears in the initial stage of the coupling process, due to the strong thermochemical effects on the TPS surface.

Apparatus And Method For Reducing Drag Of A Bluff Body In Ground Effect Using Counter-Rotating Vortex Pairs

DOEpatents

Ortega, Jason M.; Sabari, Kambiz

2005-12-27

An aerodynamic base drag reduction apparatus and method for bluff bodies, such as tractor-trailer trucks, utilizing a pair of lift surfaces extending to lift surface tips and located alongside the bluff body such as on opposing left and right side surfaces. In a flowstream substantially parallel to the longitudinal centerline of the bluff body, the pair of lift surfaces generate a pair of counter-rotating trailing vortices which confluence together in the wake of the bluff body in a direction orthogonal to the flowstream. The confluence draws or otherwise turns the flowstream, such as the flowstream passing over a top surface of the bluff body, in and around behind a trailing end of the bluff body to raise the pressure on a base surface at the trailing end and thereby reduce the aerodynamic base drag.

Apparatus And Method For Reducing Drag Of A Bluff Body In Ground Effect Using Counter-Rotating Vortex Pairs

DOEpatents

Ortega, Jason M.; Salari, Kambiz

2005-08-09

An aerodynamic base drag reduction apparatus and method for bluff bodies, such as tractor-trailer trucks, utilizing a pair of lift surfaces extending to lift surface tips and located alongside the bluff body such as on opposing left and right side surfaces. In a flowstream substantially parallel to the longitudinal centerline of the bluff body, the pair of lift surfaces generate a pair of counter-rotating trailing vortices which confluence together in the wake of the bluff body in a direction orthogonal to the flowstream. The confluence draws or otherwise turns the flowstream, such as the flowstream passing over a top surface of the bluff body, in and around behind a trailing end of the bluff body to raise the pressure on a base surface at the trailing end and thereby reduce the aerodynamic base drag.

Surface roughness effects on the subsonic aerodynamics of the Rockwell International 089B-139B orbiter

NASA Technical Reports Server (NTRS)

Ware, G. M.; Spencer, B., Jr.

1973-01-01

An experimental test program was conducted to determine the effects of vehicle surface roughness on the subsonic aerodynamic characteristics of a 0.01875 scale model of a Rockwell International Space Shuttle Configuration. Surface roughness was simulated by applying a sparce coating of carborundum grit to the complete model. Various grit sizes were investigated. Tests were conducted in the Langley Low Turbulence Pressure Tunnel at a constant nominal Mach number of 0.25 with Reynolds number varying from 2 to 12 x 10 to the 6th power per foot. Angle of attack was varied from about -2 to 28 deg at 0 deg and 6 deg angle of sideslip.

Modelling Aerodynamically Generated Sound: Recent Advances in Rotor Noise Prediction

NASA Technical Reports Server (NTRS)

Brentner, Kenneth S.

2000-01-01

A great deal of progress has been made in the modeling of aerodynamically generated sound for rotors over the past decade. The Ffowcs Williams-Hawkings (FW-H ) equation has been the foundation for much of the development. Both subsonic and supersonic quadrupole noise formulations have been developed for the prediction of high-speed impulsive noise. In an effort to eliminate the need to compute the quadrupole contribution, the FW-H has also been utilized on permeable surfaces surrounding all physical noise sources. Comparison of the Kirchhoff formulation for moving surfaces with the FW-H equation have shown that the Kirchhoff formulation for moving surfaces can give erroneous results for aeroacoustic problems.

The use of interpractive graphic displays for interpretation of surface design parameters

NASA Technical Reports Server (NTRS)

Talcott, N. A., Jr.

1981-01-01

An interactive computer graphics technique known as the Graphic Display Data method has been developed to provide a convenient means for rapidly interpreting large amounts of surface design data. The display technique should prove valuable in such disciplines as aerodynamic analysis, structural analysis, and experimental data analysis. To demonstrate the system's features, an example is presented of the Graphic Data Display method used as an interpretive tool for radiation equilibrium temperature distributions over the surface of an aerodynamic vehicle. Color graphic displays were also examined as a logical extension of the technique to improve its clarity and to allow the presentation of greater detail in a single display.

Orion Entry Performance-Based Center-of-Gravity Box

NASA Technical Reports Server (NTRS)

Rea, Jeremy R.

2010-01-01

The Orion capsule has many performance requirements for its atmospheric entry trajectory. Requirements on landing accuracy, maximum heating rate, total heat load, propellant usage, and sensed acceleration must all be satised. It is desired to define a methodology to translate the many performance requirements for an atmospheric entry trajectory into language easily understood by vehicle designers in terms of an allowable center-of-gravity box. This is possible by noting that most entry performance parameters for a capsule vehicle are mainly determined by the lift-to-drag ratio of the vehicle. However, the lift-to- drag ratio should be considered a probabilistic quantity rather than deterministic, where variations in the lift-to-drag are caused by both aerodynamic and center-of-gravity un- certainties. This paper discusses the technique used by the Orion program to define the allowable dispersions in center-of-gravity to achieve the desired entry performance while accounting for aerodynamic uncertainty.

Assessment of the Reconstructed Aerodynamics of the Mars Science Laboratory Entry Vehicle

NASA Technical Reports Server (NTRS)

Schoenenberger, Mark; Van Norman, John W.; Dyakonov, Artem A.; Karlgaard, Christopher D.; Way, David W.; Kutty, Prasad

2013-01-01

On August 5, 2012, the Mars Science Laboratory entry vehicle successfully entered Mars atmosphere, flying a guided entry until parachute deploy. The Curiosity rover landed safely in Gale crater upon completion of the Entry Descent and Landing sequence. This paper compares the aerodynamics of the entry capsule extracted from onboard flight data, including Inertial Measurement Unit (IMU) accelerometer and rate gyro information, and heatshield surface pressure measurements. From the onboard data, static force and moment data has been extracted. This data is compared to preflight predictions. The information collected by MSL represents the most complete set of information collected during Mars entry to date. It allows the separation of aerodynamic performance from atmospheric conditions. The comparisons show the MSL aerodynamic characteristics have been identified and resolved to an accuracy better than the aerodynamic database uncertainties used in preflight simulations. A number of small anomalies have been identified and are discussed. This data will help revise aerodynamic databases for future missions and will guide computational fluid dynamics (CFD) development to improved prediction codes.

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.

Supersonic Flight Dynamics Test 2: Trajectory, Atmosphere, and Aerodynamics Reconstruction

NASA Technical Reports Server (NTRS)

Karlgaard, Christopher D.; O'Farrell, Clara; Ginn, Jason M.; Van Norman, John W.

2016-01-01

The Supersonic Flight Dynamics Test is a full-scale flight test of aerodynamic decelerator technologies developed by the Low Density Supersonic Decelerator technology demonstration project. The purpose of the project is to develop and mature aerodynamic decelerator technologies for landing large-mass payloads on the surface of Mars. The technologies include a Supersonic Inflatable Aerodynamic Decelerator and supersonic parachutes. The first Supersonic Flight Dynamics Test occurred on June 28th, 2014 at the Pacific Missile Range Facility. The purpose of this test was to validate the test architecture for future tests. The flight was a success and, in addition, was able to acquire data on the aerodynamic performance of the supersonic inflatable decelerator. The Supersonic Disksail parachute developed a tear during deployment. The second flight test occurred on June 8th, 2015, and incorporated a Supersonic Ringsail parachute which was redesigned based on data from the first flight. Again, the inflatable decelerator functioned as predicted but the parachute was damaged during deployment. This paper describes the instrumentation, analysis techniques, and acquired flight test data utilized to reconstruct the vehicle trajectory, main motor thrust, atmosphere, and aerodynamics.

Data Point Averaging for Computational Fluid Dynamics Data

NASA Technical Reports Server (NTRS)

Norman, Jr., David (Inventor)

2016-01-01

A system and method for generating fluid flow parameter data for use in aerodynamic heating analysis. Computational fluid dynamics data is generated for a number of points in an area on a surface to be analyzed. Sub-areas corresponding to areas of the surface for which an aerodynamic heating analysis is to be performed are identified. A computer system automatically determines a sub-set of the number of points corresponding to each of the number of sub-areas and determines a value for each of the number of sub-areas using the data for the sub-set of points corresponding to each of the number of sub-areas. The value is determined as an average of the data for the sub-set of points corresponding to each of the number of sub-areas. The resulting parameter values then may be used to perform an aerodynamic heating analysis.

Data Point Averaging for Computational Fluid Dynamics Data

NASA Technical Reports Server (NTRS)

Norman, David, Jr. (Inventor)

2014-01-01

A system and method for generating fluid flow parameter data for use in aerodynamic heating analysis. Computational fluid dynamics data is generated for a number of points in an area on a surface to be analyzed. Sub-areas corresponding to areas of the surface for which an aerodynamic heating analysis is to be performed are identified. A computer system automatically determines a sub-set of the number of points corresponding to each of the number of sub-areas and determines a value for each of the number of sub-areas using the data for the sub-set of points corresponding to each of the number of sub-areas. The value is determined as an average of the data for the sub-set of points corresponding to each of the number of sub-areas. The resulting parameter values then may be used to perform an aerodynamic heating analysis.

Experimental Investigation of the Low-Speed Aerodynamic Characteristics of a 5.8-Percent Scale Hybrid Wing Body Configuration

NASA Technical Reports Server (NTRS)

Gatlin, Gregory M.; Vicroy, Dan D.; Carter, Melissa B.

2012-01-01

A low-speed experimental investigation has been conducted on a 5.8-percent scale Hybrid Wing Body configuration in the NASA Langley 14- by 22-Foot Subsonic Tunnel. This Hybrid Wing Body (HWB) configuration was designed with specific intention to support the NASA Environmentally Responsible Aviation (ERA) Project goals of reduced noise, emissions, and fuel burn. This HWB configuration incorporates twin, podded nacelles mounted on the vehicle upper surface between twin vertical tails. Low-speed aerodynamic characteristics were assessed through the acquisition of force and moment, surface pressure, and flow visualization data. Longitudinal and lateral-directional characteristics were investigated on this multi-component model. The effects of a drooped leading edge, longitudinal flow-through nacelle location, vertical tail shape and position, elevon deflection, and rudder deflection have been studied. The basic configuration aerodynamics, as well as the effects of these configuration variations, are presented in this paper.

Acoustical modeling study of the open test section of the NASA Langley V/STOL wind tunnel

NASA Technical Reports Server (NTRS)

Ver, I. L.; Andersen, D. W.; Bliss, D. B.

1975-01-01

An acoustic model study was carried out to identify effective sound absorbing treatment of strategically located surfaces in an open wind tunnel test section. Also an aerodynamic study done concurrently, sought to find measures to control low frequency jet pulsations which occur when the tunnel is operated in its open test section configuration. The acoustical modeling study indicated that lining of the raised ceiling and the test section floor immediately below it, results in a substantial improvement. The aerodynamic model study indicated that: (1) the low frequency jet pulsations are most likely caused or maintained by coupling of aerodynamic and aeroacoustic phenomena in the closed tunnel circuit, (2) replacing the hard collector cowl with a geometrically identical but porous fiber metal surface of 100 mks rayls flow resistance does not result in any noticable reduction of the test section noise caused by the impingement of the turbulent flow on the cowl.

A Multigrid Approach to Embedded-Grid Solvers

DTIC Science & Technology

1992-08-01

previously as a Master’s Thesis at the University of Florida. Not edited by TESCO , Inc. 12a. DISTRIBUTION / AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE...domain decomposition techniques in order to accurately model the aerodynamics of complex geometries 𔃾, 5, 11, 12, 13, 24’. Although these high...quantities subscripted by oc denote reference values in the undisturbed gas. Uv v, e e P - (10) Where • = (7b,/•)1/2, is the speed of sound in the

Is the State of the Air-Sea Interface a Factor in Rapid Intensification and Rapid Decline of Tropical Cyclones?

NASA Astrophysics Data System (ADS)

Soloviev, Alexander V.; Lukas, Roger; Donelan, Mark A.; Haus, Brian K.; Ginis, Isaac

2017-12-01

Tropical storm intensity prediction remains a challenge in tropical meteorology. Some tropical storms undergo dramatic rapid intensification and rapid decline. Hurricane researchers have considered particular ambient environmental conditions including the ocean thermal and salinity structure and internal vortex dynamics (e.g., eyewall replacement cycle, hot towers) as factors creating favorable conditions for rapid intensification. At this point, however, it is not exactly known to what extent the state of the sea surface controls tropical cyclone dynamics. Theoretical considerations, laboratory experiments, and numerical simulations suggest that the air-sea interface under tropical cyclones is subject to the Kelvin-Helmholtz type instability. Ejection of large quantities of spray particles due to this instability can produce a two-phase environment, which can attenuate gravity-capillary waves and alter the air-sea coupling. The unified parameterization of waveform and two-phase drag based on the physics of the air-sea interface shows the increase of the aerodynamic drag coefficient Cd with wind speed up to hurricane force (U10≈35 m s-1). Remarkably, there is a local Cd minimum—"an aerodynamic drag well"—at around U10≈60 m s-1. The negative slope of the Cd dependence on wind-speed between approximately 35 and 60 m s-1 favors rapid storm intensification. In contrast, the positive slope of Cd wind-speed dependence above 60 m s-1 is favorable for a rapid storm decline of the most powerful storms. In fact, the storms that intensify to Category 5 usually rapidly weaken afterward.

Effects of canard location on the aerodynamic characteristics of a blunt-nosed missile at Mach numbers of 1.5 and 2.0. [in the Ames 6x6 wind tunnel

NASA Technical Reports Server (NTRS)

Kassner, D. L.; Wettlaufer, B.

1977-01-01

A blunt-nosed missile model with nose-mounted canards and cruciform tail surfaces was tested in the Ames 6 by 6-Foot Wind Tunnel to determine the contributions of the component aerodynamic surfaces to the static aerodynamic characteristics at Mach numbers of 1.5 and 2.0 and Reynolds number of 1 million based on body diameter. Data were obtained at angles of attack ranging from -3 deg to 12 deg and canard-deflection angles from -3 deg to 15 deg for various stages of model build-up (i.e., with and without canard and/or tail surfaces). Results were obtained with the canards at two different nose locations. For the canard and tail arrangements investigated, the model was trimmable at angles of attack up to about 4 deg or 5 deg with canard deflections of 9 deg. For this blunt-nosed model, there was little effect of canard location on trim angle of attack. The tail arrangements studied provided ample pitch stability.

Shuttle ascent and shock impingement aerodynamic heating studies

NASA Technical Reports Server (NTRS)

Lanning, W. D.; Hung, F. T.

1971-01-01

The collection and analysis of aerodynamic heating data obtained from shock impingement experimental investigation were completed. The data were categorized into four interference areas; fin leading edge, wing/fuselage fin/plate corners, and space shuttle configurations. The effects of shock impingement were found to increase the heating rates 10 to 40 times the undisturbed values. A test program was completed at NASA/Langley Research Center to investigate the magnitudes and surface patterns of the mated shock interference flowfield. A 0.0065 scale thin-skin model of the MDAC 256-20 space shuttle booster mated with a Stycast model of the MDAC Internal tank orbiter was tested in the 20-inch M=6 tunnel, the 31-inch M=10 tunnel, and the 48-inch Unitary Plan Tunnel. The gap region of the ascent configuration was the principal area of interest where both thermocouple and phase-change paint data were obtained. Pressure and heat transfer distributions data on the leeward surface of a 75-degree sweep slab delta wing are presented. The effects of surface roughness on boundary layer transition and aerodynamic heating were investigated.

Long-duration heat load measurement approach by novel apparatus design and highly efficient algorithm

NASA Astrophysics Data System (ADS)

Zhu, Yanwei; Yi, Fajun; Meng, Songhe; Zhuo, Lijun; Pan, Weizhen

2017-11-01

Improving the surface heat load measurement technique for vehicles in aerodynamic heating environments is imperative, regarding aspects of both the apparatus design and identification efficiency. A simple novel apparatus is designed for heat load identification, taking into account the lessons learned from several aerodynamic heating measurement devices. An inverse finite difference scheme (invFDM) for the apparatus is studied to identify its surface heat flux from the interior temperature measurements with high efficiency. A weighted piecewise regression filter is also proposed for temperature measurement prefiltering. Preliminary verification of the invFDM scheme and the filter is accomplished via numerical simulation experiments. Three specific pieces of apparatus have been concretely designed and fabricated using different sensing materials. The aerodynamic heating process is simulated by an inductively coupled plasma wind tunnel facility. The identification of surface temperature and heat flux from the temperature measurements is performed by invFDM. The results validate the high efficiency, reliability and feasibility of heat load measurements with different heat flux levels utilizing the designed apparatus and proposed method.

Aircraft High-Lift Aerodynamic Analysis Using a Surface-Vorticity Solver

NASA Technical Reports Server (NTRS)

Olson, Erik D.; Albertson, Cindy W.

2016-01-01

This study extends an existing semi-empirical approach to high-lift analysis by examining its effectiveness for use with a three-dimensional aerodynamic analysis method. The aircraft high-lift geometry is modeled in Vehicle Sketch Pad (OpenVSP) using a newly-developed set of techniques for building a three-dimensional model of the high-lift geometry, and for controlling flap deflections using scripted parameter linking. Analysis of the low-speed aerodynamics is performed in FlightStream, a novel surface-vorticity solver that is expected to be substantially more robust and stable compared to pressure-based potential-flow solvers and less sensitive to surface perturbations. The calculated lift curve and drag polar are modified by an empirical lift-effectiveness factor that takes into account the effects of viscosity that are not captured in the potential-flow solution. Analysis results are validated against wind-tunnel data for The Energy-Efficient Transport AR12 low-speed wind-tunnel model, a 12-foot, full-span aircraft configuration with a supercritical wing, full-span slats, and part-span double-slotted flaps.

Design, Analysis and Qualification of Elevon for Reusable Launch Vehicle

NASA Astrophysics Data System (ADS)

Tiwari, S. B.; Suresh, R.; Krishnadasan, C. K.

2017-12-01

Reusable launch vehicle technology demonstrator is configured as a winged body vehicle, designed to fly in hypersonic, supersonic and subsonic regimes. The vehicle will be boosted to hypersonic speeds after which the winged body separates and descends using aerodynamic control. The aerodynamic control is achieved using the control surfaces mainly the rudder and the elevon. Elevons are deflected for pitch and roll control of the vehicle at various flight conditions. Elevons are subjected to aerodynamic, thermal and inertial loads during the flight. This paper gives details about the configuration, design, qualification and flight validation of elevon for Reusable Launch Vehicle.

Aerodynamic load distributions at transonic speeds for a close-coupled wing-canard configuration: Tabulated pressure data

NASA Technical Reports Server (NTRS)

Washburn, K. E.; Gloss, B. B.

1978-01-01

Wind tunnel studies are reported on both the canard and wing surfaces of a model that is geometrically identical to one used in several force and moment tests to provide insight into the various aerodynamic interference effects. In addition to detailed pressures measurements, the pressures were integrated to illustrate the effects of Mach number, canard location, and canard-wing interference on various aerodynamic parameters. Transonic pressure tunnel Mach numbers ranged from 0.70 to 1.20 for data taken from 0 deg to approximately 16 deg angle-of-attack at 0 deg sideslip.

Application of radiometric surface temperature for surface energy balance estimation: John Monteith's contributions

USDA-ARS?s Scientific Manuscript database

Over 25 years ago, Huband and Monteith paper’s investigating the radiative surface temperature and the surface energy balance of a wheat canopy, highlighted the key issues in computing fluxes with radiometric surface temperature. These included the relationship between radiometric and aerodynamic s...

Effect of wind tunnel acoustic modes on linear oscillating cascade aerodynamics

NASA Technical Reports Server (NTRS)

Buffum, Daniel H.; Fleeter, Sanford

1993-01-01

The aerodynamics of a biconvex airfoil cascade oscillating in torsion is investigated using the unsteady aerodynamic influence coefficient technique. For subsonic flow and reduced frequencies as large as 0.9, airfoil surface unsteady pressures resulting from oscillation of one of the airfoils are measured using flush-mounted high-frequency-response pressure transducers. The influence coefficient data are examined in detail and then used to predict the unsteady aerodynamics of a cascade oscillating at various interblade phase angles. These results are correlated with experimental data obtained in the traveling-wave mode of oscillation and linearized analysis predictions. It is found that the unsteady pressure disturbances created by an oscillating airfoil excite wind tunnel acoustic modes which have detrimental effects on the experimental data. Acoustic treatment is proposed to rectify this problem.

Aerodynamic Characteristics of a Revised Target Drone Vehicle at Mach Numbers from 1.60 to 2.86

NASA Technical Reports Server (NTRS)

Blair, A. B., Jr.; Babb, C. Donald

1968-01-01

An investigation has been conducted in the Langley Unitary Plan wind tunnel to determine the aerodynamic characteristics of a revised target drone vehicle through a Mach number range from 1.60 to 2.86. The vehicle had canard surfaces and a swept clipped-delta wing with twin tip-mounted vertical tails.

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.

The influence of sweep on the aerodynamic loading of an oscillating NACA 0012 airfoil. Volume 1: Technical report

NASA Technical Reports Server (NTRS)

St.hilaire, A. O.; Carta, F. O.; Fink, M. R.; Jepson, W. D.

1979-01-01

Aerodynamic experiments were performed on an oscillating NACA 0012 airfoil utilizing a tunnel-spanning wing in both unswept and 30 degree swept configurations. The airfoil was tested in steady state and in oscillatory pitch about the quarter chord. The unsteady aerodynamic loading was measured using pressure transducers along the chord. Numerical integrations of the unsteady pressure transducer responses were used to compute the normal force, chord force, and moment components of the induced loading. The effects of sweep on the induced aerodynamic load response was examined. For the range of parameters tested, it was found that sweeping the airfoil tends to delay the onset of dynamic stall. Sweeping was also found to reduce the magnitude of the unsteady load variation about the mean response. It was determined that at mean incidence angles greater than 9 degrees, sweep tends to reduce the stability margin of the NACA 0012 airfoil; however, for all cases tested, the airfoil was found to be stable in pure pitch. Turbulent eddies were found to convect downstream above the upper surface and generate forward-moving acoustic waves at the trailing edge which move upstream along the lower surface.

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.

Improving the Unsteady Aerodynamic Performance of Transonic Turbines using Neural Networks

NASA Technical Reports Server (NTRS)

Rai, Man Mohan; Madavan, Nateri K.; Huber, Frank W.

1999-01-01

A recently developed neural net-based aerodynamic design procedure is used in the redesign of a transonic turbine stage to improve its unsteady aerodynamic performance. The redesign procedure used incorporates the advantages of both traditional response surface methodology and neural networks by employing a strategy called parameter-based partitioning of the design space. Starting from the reference design, a sequence of response surfaces based on both neural networks and polynomial fits are constructed to traverse the design space in search of an optimal solution that exhibits improved unsteady performance. The procedure combines the power of neural networks and the economy of low-order polynomials (in terms of number of simulations required and network training requirements). A time-accurate, two-dimensional, Navier-Stokes solver is used to evaluate the various intermediate designs and provide inputs to the optimization procedure. The procedure yielded a modified design that improves the aerodynamic performance through small changes to the reference design geometry. These results demonstrate the capabilities of the neural net-based design procedure, and also show the advantages of including high-fidelity unsteady simulations that capture the relevant flow physics in the design optimization process.

Neural Net-Based Redesign of Transonic Turbines for Improved Unsteady Aerodynamic Performance

NASA Technical Reports Server (NTRS)

Madavan, Nateri K.; Rai, Man Mohan; Huber, Frank W.

1998-01-01

A recently developed neural net-based aerodynamic design procedure is used in the redesign of a transonic turbine stage to improve its unsteady aerodynamic performance. The redesign procedure used incorporates the advantages of both traditional response surface methodology (RSM) and neural networks by employing a strategy called parameter-based partitioning of the design space. Starting from the reference design, a sequence of response surfaces based on both neural networks and polynomial fits are constructed to traverse the design space in search of an optimal solution that exhibits improved unsteady performance. The procedure combines the power of neural networks and the economy of low-order polynomials (in terms of number of simulations required and network training requirements). A time-accurate, two-dimensional, Navier-Stokes solver is used to evaluate the various intermediate designs and provide inputs to the optimization procedure. The optimization procedure yields a modified design that improves the aerodynamic performance through small changes to the reference design geometry. The computed results demonstrate the capabilities of the neural net-based design procedure, and also show the tremendous advantages that can be gained by including high-fidelity unsteady simulations that capture the relevant flow physics in the design optimization process.

Model Reduction of Computational Aerothermodynamics for Multi-Discipline Analysis in High Speed Flows

NASA Astrophysics Data System (ADS)

Crowell, Andrew Rippetoe

This dissertation describes model reduction techniques for the computation of aerodynamic heat flux and pressure loads for multi-disciplinary analysis of hypersonic vehicles. NASA and the Department of Defense have expressed renewed interest in the development of responsive, reusable hypersonic cruise vehicles capable of sustained high-speed flight and access to space. However, an extensive set of technical challenges have obstructed the development of such vehicles. These technical challenges are partially due to both the inability to accurately test scaled vehicles in wind tunnels and to the time intensive nature of high-fidelity computational modeling, particularly for the fluid using Computational Fluid Dynamics (CFD). The aim of this dissertation is to develop efficient and accurate models for the aerodynamic heat flux and pressure loads to replace the need for computationally expensive, high-fidelity CFD during coupled analysis. Furthermore, aerodynamic heating and pressure loads are systematically evaluated for a number of different operating conditions, including: simple two-dimensional flow over flat surfaces up to three-dimensional flows over deformed surfaces with shock-shock interaction and shock-boundary layer interaction. An additional focus of this dissertation is on the implementation and computation of results using the developed aerodynamic heating and pressure models in complex fluid-thermal-structural simulations. Model reduction is achieved using a two-pronged approach. One prong focuses on developing analytical corrections to isothermal, steady-state CFD flow solutions in order to capture flow effects associated with transient spatially-varying surface temperatures and surface pressures (e.g., surface deformation, surface vibration, shock impingements, etc.). The second prong is focused on minimizing the computational expense of computing the steady-state CFD solutions by developing an efficient surrogate CFD model. The developed two-pronged approach is found to exhibit balanced performance in terms of accuracy and computational expense, relative to several existing approaches. This approach enables CFD-based loads to be implemented into long duration fluid-thermal-structural simulations.

Characteristics of an Airfoil as Affected by Fabric Sag

NASA Technical Reports Server (NTRS)

Ward, Kenneth E

1932-01-01

This report presents the results of tests made at a high value of the Reynolds Number in the N.A.C.A. variable-density wind tunnel to determine the aerodynamic characteristics of an airfoil as affected by fabric sag. Tests were made of two Gottingen 387 airfoils, one having the usual smooth surface and the other having a surface modified to simulate two types of fabric sag. The results of these tests indicate that the usual sagging of the wind covering between ribs has a very small effect on the aerodynamic characteristics of an airfoil.

Relationships between topographic roughness and aeolian processes

NASA Technical Reports Server (NTRS)

Greeley, Ronald; Lancaster, N.; Gaddis, L.; Rasmussen, K. R.; White, B. R.; Saunders, R. S.; Wall, S.; Dobrovolskis, Anthony R.; Iversen, J. D.

1991-01-01

The interaction between winds and desert surfaces has important implications for sediment transport on Earth, Mars, and Venus, and for understanding the relationships between radar backscatter and aerodynamic roughness as part of the NASA Shuttle Imaging radar (SIR-C) Mission. Here, researchers report results from measurements of boundary layer wind profiles and surface roughness at sites in Death Valley and discuss their implications. The sites included a flat to undulating gravel and sand reg, alluvial fans, and a playa. Estimates of average particle size composition of Death Valley sites and arithmetic mean values of aerodynamic roughness are given in tabular form.

Relating Vegetation Aerodynamic Roughness Length to Interferometric SAR Measurements

NASA Technical Reports Server (NTRS)

Saatchi, Sassan; Rodriquez, Ernesto

1998-01-01

In this paper, we investigate the feasibility of estimating aerodynamic roughness parameter from interferometric SAR (INSAR) measurements. The relation between the interferometric correlation and the rms height of the surface is presented analytically. Model simulations performed over realistic canopy parameters obtained from field measurements in boreal forest environment demonstrate the capability of the INSAR measurements for estimating and mapping surface roughness lengths over forests and/or other vegetation types. The procedure for estimating this parameter over boreal forests using the INSAR data is discussed and the possibility of extending the methodology over tropical forests is examined.

Computer program for supersonic Kernel-function flutter analysis of thin lifting surfaces

NASA Technical Reports Server (NTRS)

Cunningham, H. J.

1974-01-01

This report describes a computer program (program D2180) that has been prepared to implement the analysis described in (N71-10866) for calculating the aerodynamic forces on a class of harmonically oscillating planar lifting surfaces in supersonic potential flow. The planforms treated are the delta and modified-delta (arrowhead) planforms with subsonic leading and supersonic trailing edges, and (essentially) pointed tips. The resulting aerodynamic forces are applied in a Galerkin modal flutter analysis. The required input data are the flow and planform parameters including deflection-mode data, modal frequencies, and generalized masses.

Experimental and theoretical study of aerodynamic characteristics of some lifting bodies at angles of attack from -10 degrees to 53 degrees at Mach numbers from 2.30 to 4.62

NASA Technical Reports Server (NTRS)

Spearman, M. Leroy; Torres, Abel O.

1994-01-01

Lifting bodies are of interest for possible use as space transportation vehicles because they have the volume required for significant payloads and the aerodynamic capability to negotiate the transition from high angles of attack to lower angles of attack (for cruise flight) and thus safely reenter the atmosphere and perform conventional horizontal landings. Results are presented for an experimental and theoretical study of the aerodynamic characteristics at supersonic speeds for a series of lifting bodies with 75 deg delta planforms, rounded noses, and various upper and lower surface cambers. The camber shapes varied in thickness and in maximum thickness location, and hence in body volume. The experimental results were obtained in the Langley Unitary Plan Wind Tunnel for both the longitudinal and the lateral aerodynamic characteristics. Selected experimental results are compared with calculated results obtained through the use of the Hypersonic Arbitrary-Body Aerodynamic Computer Program.

Investigation of oscillating cascade aerodynamics by an experimental influence coefficient technique

NASA Technical Reports Server (NTRS)

Buffum, Daniel H.; Fleeter, Sanford

1988-01-01

Fundamental experiments are performed in the NASA Lewis Transonic Oscillating Cascade Facility to investigate the torsion mode unsteady aerodynamics of a biconvex airfoil cascade at realistic values of the reduced frequency for all interblade phase angles at a specified mean flow condition. In particular, an unsteady aerodynamic influence coefficient technique is developed and utilized in which only one airfoil in the cascade is oscillated at a time and the resulting airfoil surface unsteady pressure distribution measured on one dynamically instrumented airfoil. The unsteady aerodynamics of an equivalent cascade with all airfoils oscillating at a specified interblade phase angle are then determined through a vector summation of these data. These influence coefficient determined oscillation cascade data are correlated with data obtained in this cascade with all airfoils oscillating at several interblade phase angle values. The influence coefficients are then utilized to determine the unsteady aerodynamics of the cascade for all interblade phase angles, with these unique data subsequently correlated with predictions from a linearized unsteady cascade model.

Study on the Influence of the Convoy Rolling over Aerodynamic Resistance

NASA Astrophysics Data System (ADS)

Iozsa, D.; Stan, C.; Ilea, L.

2017-10-01

The aim of the study is to investigate how the aerodynamic resistance is influenced by the convoy rolling and to see how much this is possible by varying the distance between trucks. Then to see how the gains correlate with the position occupied by the truck in the convoy. The study starts from current research on the premises of running in convoy. Aerodynamic analysis was performed using software finite element of Computational Fluid Dynamics (CFD) type, where it was modeled the convoy rolling of a variable number of trucks. The number of trucks and the distance between them was varied in the model in order to acquire an understanding of the flow field around the trucks and how the distance between them can improve the aerodynamic parameters. The results are presented in the form of streamlines of the air, which indicates the air volume travel speed and direction and of the pressure distribution on the surface of the body. The most significant drop in pressure on the front surface was obtained for the second truck of the convoy, whereas for the following ones the reduction was less important. The participation in a convoy of more than two trucks is justified by the reduction of the whirls that appear and by the uniform air flow. The main advantage of running in convoy mode is to decrease aerodynamic resistance, with beneficial consequences on economic and ecological parameters. Continuing work from here on, it could be analyzed the impact of changing the distance between trucks on the aerodynamic coefficient. The results of CFD simulations need to be verified with experimental data, such as wind-tunnel test, to ensure reliability of the results.

Aerodynamic method for obtaining the soil water retention curve

NASA Astrophysics Data System (ADS)

Alekseev, V. V.; Maksimov, I. I.

2013-07-01

A new method for the rapid plotting of the soil water retention curve (SWRC) has been proposed that considers the soil water as an environment limited by the soil solid phase on one side and by the soil air on the other side. Both contact surfaces have surface energies, which play the main role in water retention. The use of an idealized soil model with consideration for the nonequilibrium thermodynamic laws and the aerodynamic similarity principles allows us to estimate the volumetric specific surface areas of soils and, using the proposed pedotransfer function (PTF), to plot the SWRC. The volumetric specific surface area of the solid phase, the porosity, and the specific free surface energy at the water-air interface are used as the SWRC parameters. Devices for measuring the parameters are briefly described. The differences between the proposed PTF and the experimental data have been analyzed using the statistical processing of the data.

Discussion of flight experiments with an entry research vehicle

NASA Technical Reports Server (NTRS)

Potter, J. L.

1985-01-01

The focus of interest is the maneuvering flight of advanced entry vehicles operating at altitudes above 50 km and at velocities of 5 to 8 km/s. Information resulting in more accurate aerodynamic analysis is sought and measurement techniques that appear to be applicable are identified. Measurements discussed include: shock layer or boundary layer profiles of velocity, temperature, species mass fractions, and other gas properties associated with aerodynamic heating; surface energy transfer process; nonequilibrium flow processes and pressure distribution; separated, vortic leeside flow of nonequilibrium fluid; boundary layer transition on highly swept configurations; and shock and surface slip and gas/surface interaction. Further study should focus on evolving measurement techniques, installation requirements, and on identification of the portions of flights where successful results seem probable.

Influence of airfoil geometry on delta wing leading-edge vortices and vortex-induced aerodynamics at supersonic speeds

NASA Technical Reports Server (NTRS)

Wood, Richard M.; Byrd, James E.; Wesselmann, Gary F.

1992-01-01

An assessment of the influence of airfoil geometry on delta wing leading edge vortex flow and vortex induced aerodynamics at supersonic speeds is discussed. A series of delta wing wind tunnel models were tested over a Mach number range from 1.7 to 2.0. The model geometric variables included leading edge sweep and airfoil shape. Surface pressure data, vapor screen, and oil flow photograph data were taken to evaluate the complex structure of the vortices and shocks on the family of wings tested. The data show that airfoil shape has a significant impact on the wing upper surface flow structure and pressure distribution, but has a minimal impact on the integrated upper surface pressure increments.

Geometry of the hemispherical radiometric footprint over plant canopies

NASA Astrophysics Data System (ADS)

Marcolla, B.; Cescatti, A.

2017-11-01

Radiometric measurements of hemispherical surface reflectance and long-wave irradiance are required to quantify the broadband albedo and the outgoing thermal radiation. These observations are typically integrated with eddy covariance measurements of sensible and latent heat fluxes to characterize the surface energy budget. While the aerodynamic footprint has been widely investigated, the geometry of the hemispherical radiometric footprint over plant canopies has been rarely tackled. In the present work, the size and shape of the hemispherical radiometric footprint are formalized for a bare surface and in presence of a vegetation cover. For this purpose, four idealized canopies are analyzed and the dependency of the radiometric footprint on leaf area index and canopy height is explored. Besides, the radiometric footprint is compared with the aerodynamic footprint in conditions of neutral stability. It was observed that almost 100% of the hemispherical radiometric signal originates within a distance of a few radiometer heights, while only about 50-80% of the cumulative aerodynamic signal is generated within a distance of about 20 sensor heights. In order to achieve comparable extensions of the footprint areas, hemispherical radiometric measurements should therefore be taken about 6-15 times higher than turbulent flux ones, depending on the vegetation type. The analysis also highlights that the size of the radiative footprint decreases at increasing leaf area index, whereas the aerodynamic footprint shows an opposite behavior. For the abovementioned reasons, this work may support the interpretation of energy flux measurements and the optimal design of eddy covariance stations located in heterogeneous sites.

Aerodynamic Influence of Added Surfaces on the Performance Characteristics of a Sports Car

NASA Astrophysics Data System (ADS)

Thangadurai, Murugan; Kumar, Rajesh; Rana, Subhas Chandra; Chatterjee, Dipankar

2018-05-01

External aerodynamics plays a vital role in designing high-speed vehicles since a reduction in drag and positive lift generation are principal concerns in vehicle aerodynamics to ensure superior performance, comfort, and vehicle stability. In the present study, the effect of added surfaces such as NACA 2412 wings and wedge type spoiler at the rear end of a sports car are examined in detail using three-dimensional numerical simulations substantiated with lab scale experiments. The simulations are performed by solving Reynolds-averaged Navier-Stokes equations with a realizable k-ɛ turbulence model using ANSYS Fluent software for Reynolds numbers 9.1 × 106, 1.37 × 107 and 1.82 × 107. The results obtained from simulations are validated with the experiments performed on a scale down model at the low-speed wind tunnel using a six component external pyramidal balance. The variation in the wake flow field of the vehicles with different added surfaces are demonstrated using pressure and velocity contours, velocity vectors at the rear end, and the turbulent kinetic energy distribution plots. It is observed that the positive lift coefficient of the base model is reduced drastically by incorporating a single wing at the rear end of the vehicle. The aerodynamics coefficients obtained from different configurations suggest that the two wing configuration has lesser drag than the wedge type spoiler though, the negative lift is higher with a wedge than the two wing configuration.

Aero-thermal optimization of film cooling flow parameters on the suction surface of a high pressure turbine blade

NASA Astrophysics Data System (ADS)

El Ayoubi, Carole; Hassan, Ibrahim; Ghaly, Wahid

2012-11-01

This paper aims to optimize film coolant flow parameters on the suction surface of a high-pressure gas turbine blade in order to obtain an optimum compromise between a superior cooling performance and a minimum aerodynamic penalty. An optimization algorithm coupled with three-dimensional Reynolds-averaged Navier Stokes analysis is used to determine the optimum film cooling configuration. The VKI blade with two staggered rows of axially oriented, conically flared, film cooling holes on its suction surface is considered. Two design variables are selected; the coolant to mainstream temperature ratio and total pressure ratio. The optimization objective consists of maximizing the spatially averaged film cooling effectiveness and minimizing the aerodynamic penalty produced by film cooling. The effect of varying the coolant flow parameters on the film cooling effectiveness and the aerodynamic loss is analyzed using an optimization method and three dimensional steady CFD simulations. The optimization process consists of a genetic algorithm and a response surface approximation of the artificial neural network type to provide low-fidelity predictions of the objective function. The CFD simulations are performed using the commercial software CFX. The numerical predictions of the aero-thermal performance is validated against a well-established experimental database.

14 CFR 25.445 - Auxiliary aerodynamic surfaces.

Code of Federal Regulations, 2011 CFR

2011-01-01

... AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Structure Control Surface and System Loads..., and gusts as specified in § 25.341(a) acting at any orientation at right angles to the flight path. (b...

14 CFR 25.445 - Auxiliary aerodynamic surfaces.

Code of Federal Regulations, 2010 CFR

2010-01-01

... AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Structure Control Surface and System Loads..., and gusts as specified in § 25.341(a) acting at any orientation at right angles to the flight path. (b...

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

Potential application of artificial concepts to aerodynamic simulation

NASA Technical Reports Server (NTRS)

Kutler, P.; Mehta, U. B.; Andrews, A.

1984-01-01

The concept of artificial intelligence as it applies to computational fluid dynamics simulation is investigated. How expert systems can be adapted to speed the numerical aerodynamic simulation process is also examined. A proposed expert grid generation system is briefly described which, given flow parameters, configuration geometry, and simulation constraints, uses knowledge about the discretization process to determine grid point coordinates, computational surface information, and zonal interface parameters.

Aerodynamic Forces and Moments of a Seaplane on the Water

NASA Technical Reports Server (NTRS)

Kohler, M

1933-01-01

This report gives the results of wind-tunnel tests with a seaplane model as a contribution to the solution of the aerodynamic problems. In the tests it was assumed that the seaplane rested motionless on the water and was exposed, in various positions with respect to the supposedly flat surface of the water, to a uniform air current 0 to 360 degrees.

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.

Noncontact surface tension and viscosity measurements of molten oxides with a pressurized hybrid electrostatic-aerodynamic levitator

NASA Astrophysics Data System (ADS)

Ishikawa, Takehiko; Yu, Jianding; Paradis, Paul-François

2006-05-01

In order to measure the surface tension and the viscosity of molten oxides, the oscillation drop technique has been applied on a pressurized hybrid electrostatic-aerodynamic levitator. To suppress the electrical discharge between the top and bottom electrodes, the drop excitation method which has been used with high vacuum electrostatic levitators has been modified. As a demonstration, the surface tension and viscosity of liquid BaTiO3 were measured using this new method. Over the 1500-2000K interval, the surface tension was measured as γ(T )=349-0.03 (T-Tm) (10-3N/m), where Tm=1893K is the melting temperature. Similarly, the viscosity was determined as η(T )=0.53exp[5.35×104/(RT)](10-3Pas) over the same temperature interval.

Effects and mechanism on Kapton film under ozone exposure in a ground near space simulator

NASA Astrophysics Data System (ADS)

Wei, Qiang; Yang, Guimin; Liu, Gang; Jiang, Haifu; Zhang, Tingting

2018-05-01

The effect on aircraft materials in the near space environment is a key part of air-and-space integration research. Ozone and aerodynamic fluids are important organizational factors in the near space environment and both have significant influences on the performance of aircraft materials. In the present paper a simulated ozone environment was used to test polyimide material that was rotated at the approximate velocity of 150-250 m/s to form an aerodynamic fluid field. The goal was to evaluate the performance evolution of materials under a comprehensive environment of ozone molecular corrosion and aerodynamic fluids. The research results show that corrosion and sputtering by ozone molecules results in Kapton films exhibiting a rugged "carpet-like" morphology exhibits an increase in surface roughness. The morphology after ozone exposure led to higher surface roughness and an increase in surface optical diffuse reflection, which is expressed by the lower optical transmittance and the gradual transition from light orange to brown. The mass loss test, XPS, and FTIR analysis show that the molecular chains on the surface of the Kapton film are destroyed resulting in Csbnd C bond breaking to form small volatile molecules such as CO2 or CO, which are responsible for a linear increase in mass loss per unit area. The Csbnd N and Csbnd O structures exhibit weakening tendency under ozone exposure. The present paper explores the evaluation method for Kapton's adaptability under the ozone exposure test in the near space environment, and elucidates the corrosion mechanism and damage mode of the polyimide material under the combined action of ozone corrosion and the aerodynamic fluid. This work provides a methodology for studying materials in the near-space environment.

Aerodynamic design of a rotor blade for minimum noise radiation

NASA Technical Reports Server (NTRS)

Karamcheti, K.; Yu, Y. H.

1974-01-01

An analysis of the aerodynamic design of a hovering rotor blade for obtaining minimum aerodynamic rotor noise has been carried out. In this analysis, which is based on both acoustical and aerodynamic considerations, attention is given only to the rotational noise due to the pressure fluctuations on the blade surfaces. The lift distribution obtained in this analysis has different characteristics from those of the conventional distribution. The present distribution shows negative lift values over a quarter of the span from the blade tip, and a maximum lift at about the midspan. Results are presented to show that the noise field is considerably affected by the shape of the lift distribution along the blade and that noise reduction of about 5 dB may be obtained by designing the rotor blade to yield minimum noise.

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.

Unsteady aerodynamic analysis of space shuttle vehicles. Part 2: Steady and unsteady aerodynamics of sharp-edged delta wings

NASA Technical Reports Server (NTRS)

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

1973-01-01

An analysis of the steady and unsteady aerodynamics of sharp-edged slender wings has been performed. The results show that slender wing theory can be modified to give the potential flow static and dynamic characteristics in incompressible flow. A semiempirical approximation is developed for the vortex-induced loads, and it is shown that the analytic approximation for sharp-edged slender wings gives good prediction of experimentally determined steady and unsteady aerodynamics at M = 0 and M = 1. The predictions are good not only for delta wings but also for so-called arrow and diamond wings. The results indicate that the effects of delta planform lifting surfaces can be included in a simple manner when determining elastic launch vehicle dynamic characteristics. For Part 1 see (N73-32763).

Space shuttle: Static aerodynamic characteristics characteristics and control effectiveness for McDonnell-Douglas orbiter configuration for Mach number range of 0.4 to 5.0

NASA Technical Reports Server (NTRS)

Ellis, R. R.

1971-01-01

An experimental aerodynamic wind tunnel investigation was conducted employing a 0.00325 scale model of the McDonnell-Douglas space shuttle orbiter configuration. This investigation was conducted in the NASA/Marshall Space Flight Center 14- by 14- inch trisonic wind tunnel. The investigation was to determine the aerodynamic characteristics of the orbiter over the Mach number range of 0.4 to 5.0, an angle of attack variation from -4 degrees to 50 degrees, and -6 degrees to 9 degrees angle of sideslip. Control surface effectiveness was investigated for elevator, aileron, and rudder deflections.

Aerodynamic preliminary analysis system 2. Part 2: User's manuals

NASA Technical Reports Server (NTRS)

Divan, P.

1981-01-01

An aerodynamic analysis system based on potential theory at subsonic/supersonic speeds and impact type finite element solutions at hypersonic conditions is described. Three dimensional configurations having multiple nonplanar surfaces of arbitrary planform and bodies of noncircular contour may be analyzed. Static, rotary, and control longitudinal and lateral directional chracteristics may be generated. The analysis has been implemented on a time sharing system in conjunction with an input tablet digitizer and an interactive graphics input/output display and editing terminal to maximize its responsiveness to the preliminary analysis problem. Typical simulation indicates that program provides an efficient analysis for systematically performing various aerodynamic configuration tradeoff and evaluation studies.

Aerodynamic preliminary analysis system 2. Part 1: Theory

NASA Technical Reports Server (NTRS)

Bonner, E.; Clever, W.; Dunn, K.

1991-01-01

An aerodynamic analysis system based on potential theory at subsonic and/or supersonic speeds and impact type finite element solutions at hypersonic conditions is described. Three dimensional configurations having multiple nonplanar surfaces of arbitrary planform and bodies of noncircular contour may be analyzed. Static, rotary, and control longitudinal and lateral directional characteristics may be generated. The analysis was implemented on a time sharing system in conjunction with an input tablet digitizer and an interactive graphics input/output display and editing terminal to maximize its responsiveness to the preliminary analysis problem. The program provides an efficient analysis for systematically performing various aerodynamic configuration tradeoff and evaluation studies.

CFD Simulations of the Supersonic Inflatable Aerodynamic Decelerator (SIAD) Ballistic Range Tests

NASA Technical Reports Server (NTRS)

Brock, Joseph; Stern, Eric; Wilder, Michael

2017-01-01

A series of ballistic range tests were performed on a scaled model of the Supersonic Flight Demonstration Test (SFDT) intended to test the Supersonic Inflatable Aerodynamic Decelerator (SIAD) geometry. The purpose of these experiments were to provide aerodynamic coefficients of the vehicle to aid in mission and vehicle design. The experimental data spans the moderate Mach number range, $3.8-2.0$, with a total angle of attack ($alpha_T$) range, $10o-20o$. These conditions are intended to span the Mach-$alpha$ space for the majority of the SFDT experiment. In an effort to validate the predictive capabilities of Computational Fluid Dynamics (CFD) for free-flight aerodynamic behavior, numerical simulations of the ballistic range experiment are performed using the unstructured finite volume Navier-Stokes solver, US3D. Comparisons to raw vehicle attitude, and post-processed aerodynamic coefficients are made between simulated results and experimental data. The resulting comparisons for both raw model attitude and derived aerodynamic coefficients show good agreement with experimental results. Additionally, near body pressure field values for each trajectory simulated are investigated. Extracted surface and wake pressure data gives further insights into dynamic flow coupling leading to a potential mechanism for dynamic instability.

Development of a linearized unsteady Euler analysis for turbomachinery blade rows

NASA Technical Reports Server (NTRS)

Verdon, Joseph M.; Montgomery, Matthew D.; Kousen, Kenneth A.

1995-01-01

A linearized unsteady aerodynamic analysis for axial-flow turbomachinery blading is described in this report. The linearization is based on the Euler equations of fluid motion and is motivated by the need for an efficient aerodynamic analysis that can be used in predicting the aeroelastic and aeroacoustic responses of blade rows. The field equations and surface conditions required for inviscid, nonlinear and linearized, unsteady aerodynamic analyses of three-dimensional flow through a single, blade row operating within a cylindrical duct, are derived. An existing numerical algorithm for determining time-accurate solutions of the nonlinear unsteady flow problem is described, and a numerical model, based upon this nonlinear flow solver, is formulated for the first-harmonic linear unsteady problem. The linearized aerodynamic and numerical models have been implemented into a first-harmonic unsteady flow code, called LINFLUX. At present this code applies only to two-dimensional flows, but an extension to three-dimensions is planned as future work. The three-dimensional aerodynamic and numerical formulations are described in this report. Numerical results for two-dimensional unsteady cascade flows, excited by prescribed blade motions and prescribed aerodynamic disturbances at inlet and exit, are also provided to illustrate the present capabilities of the LINFLUX analysis.

Aeroacoustic and aerodynamic applications of the theory of nonequilibrium thermodynamics

NASA Technical Reports Server (NTRS)

Horne, W. Clifton; Smith, Charles A.; Karamcheti, Krishnamurty

1991-01-01

Recent developments in the field of nonequilibrium thermodynamics associated with viscous flows are examined and related to developments to the understanding of specific phenomena in aerodynamics and aeroacoustics. A key element of the nonequilibrium theory is the principle of minimum entropy production rate for steady dissipative processes near equilibrium, and variational calculus is used to apply this principle to several examples of viscous flow. A review of nonequilibrium thermodynamics and its role in fluid motion are presented. Several formulations are presented of the local entropy production rate and the local energy dissipation rate, two quantities that are of central importance to the theory. These expressions and the principle of minimum entropy production rate for steady viscous flows are used to identify parallel-wall channel flow and irrotational flow as having minimally dissipative velocity distributions. Features of irrotational, steady, viscous flow near an airfoil, such as the effect of trailing-edge radius on circulation, are also found to be compatible with the minimum principle. Finally, the minimum principle is used to interpret the stability of infinitesimal and finite amplitude disturbances in an initially laminar, parallel shear flow, with results that are consistent with experiment and linearized hydrodynamic stability theory. These results suggest that a thermodynamic approach may be useful in unifying the understanding of many diverse phenomena in aerodynamics and aeroacoustics.

Analysis of Nonplanar Wing-tip-mounted Lifting Surfaces on Low-speed Airplanes

NASA Technical Reports Server (NTRS)

Vandam, C. P.; Roskam, J.

1983-01-01

Nonplanar wing tip mounted lifting surfaces reduce lift induced drag substantially. Winglets, which are small, nearly vertical, winglike surfaces, are an example of these devices. To achieve reduction in lift induced drag, winglets produce significant side forces. Consequently, these surfaces can seriously affect airplane lateral directional aerodynamic characteristics. Therefore, the effects of nonplanar wing tip mounted surfaces on the lateral directional stability and control of low speed general aviation airplanes were studied. The study consists of a theoretical and an experimental, in flight investigation. The experimental investigation involves flight tests of winglets on an agricultural airplane. Results of these tests demonstrate the significant influence of winglets on airplane lateral directional aerodynamic characteristics. It is shown that good correlations exist between experimental data and theoretically predicted results. In addition, a lifting surface method was used to perform a parametric study of the effects of various winglet parameters on lateral directional stability derivatives of general aviation type wings.

Real-Time Flight Envelope Monitoring System

NASA Technical Reports Server (NTRS)

Kerho, Michael; Bragg, Michael B.; Ansell, Phillip J.

2012-01-01

The objective of this effort was to show that real-time aircraft control-surface hinge-moment information could be used to provide a robust and reliable prediction of vehicle performance and control authority degradation. For a given airfoil section with a control surface -- be it a wing with an aileron, rudder, or elevator -- the control-surface hinge moment is sensitive to the aerodynamic characteristics of the section. As a result, changes in the aerodynamics of the section due to angle-of-attack or environmental effects such as icing, heavy rain, surface contaminants, bird strikes, or battle damage will affect the control surface hinge moment. These changes include both the magnitude of the hinge moment and its sign in a time-averaged sense, and the variation of the hinge moment with time. The current program attempts to take the real-time hinge moment information from the aircraft control surfaces and develop a system to predict aircraft envelope boundaries across a range of conditions, alerting the flight crew to reductions in aircraft controllability and flight boundaries.

Secondary flow and heat transfer control in gas turbine inlet nozzle guide vanes

NASA Astrophysics Data System (ADS)

Burd, Steven Wayne

1998-12-01

Endwall heat transfer is a very serious problem in the inlet nozzle guide vane region of gas turbine engines. To resolve heat transfer concerns and provide the desired thermal protection, modern cooling flows for the vane endwalls tend to be excessive leading to lossy and inefficient designs. Coolant introduction is further complicated by the flow patterns along vane endwall surfaces. They are three-dimensional and dominated by strong, complex secondary flows. To achieve performance goals for next-generation engines, more aerodynamically efficient and advanced cooling concepts, including combustor bleed cooling, must be investigated. To this end, the overall performance characteristics of several combustor bleed flow designs are assessed in this experimental study. In particular, their contributions toward secondary flow control and component cooling are documented. Testing is performed in a large-scale, guide vane simulator comprised of three airfoils encased between one contoured and one flat endwall. Core flow is supplied to this simulator at an inlet chord Reynolds number of 350,000 and turbulence intensity of 9.5%. Combustor bleed cooling flow is injected through the contoured endwall via inclined slots. The slots vary in cross-sectional area, have equivalent slot widths, and are positioned with their leeward edges 10% of the axial chord ahead of the airfoil leading edges. Measurements with hot-wire anemometry characterize the inlet and exit flow fields of the cascade. Total and static pressure measurements document aerodynamic performance. Thermocouple measurements detail thermal fields and permit evaluation of surface adiabatic effectiveness. To elucidate the effects of bleed injection, data are compared to an experiment taken without bleed. The influence of bleed mass flow rate and slot geometry on the aerodynamic losses and thermal protection arc given. This study suggests that such combustor bleed flow cooling offers significant thermal protection without imposing aerodynamic penalties. Such performance is contrary to the performance of present vane cooling schemes. The results of this investigation support designs which incorporate combustor coolant injection upstream of the airfoil leading edges. To complement, a short exploratory study regarding the effects of surface roughness was also performed. Results indicate modified cooling performance and significantly higher aerodynamic losses with rough surfaces.

Application of Pressure Sensitive Paint in Hypersonic Flows

NASA Technical Reports Server (NTRS)

Jules, Kenol; Carbonaro, Mario; Zemsch, Stephan

1995-01-01

It is well known in the aerodynamic field that pressure distribution measurement over the surface of an aircraft model is a problem in experimental aerodynamics. For one thing, a continuous pressure map can not be obtained with the current experimental methods since they are discrete. Therefore, interpolation or CFD methods must be used for a more complete picture of the phenomenon under study. For this study, a new technique was investigated which would provide a continuous pressure distribution over the surface under consideration. The new method is pressure sensitive paint. When pressure sensitive paint is applied to an aerodynamic surface and placed in an operating wind-tunnel under appropriate lighting, the molecules luminesce as a function of the local pressure of oxygen over the surface of interest during aerodynamic flow. The resulting image will be brightest in the areas of low pressure (low oxygen concentration), and less intense in the areas of high pressure (where oxygen is most abundant on the surface). The objective of this investigation was to use pressure sensitive paint samples from McDonnell Douglas (MDD) for calibration purpose in order to assess the response of the paint under appropriate lighting and to use the samples over a flat plate/conical fin mounted at 75 degrees from the center of the plate in order to study the shock/boundary layer interaction at Mach 6 in the Von Karman wind-tunnel. From the result obtained it was concluded that temperature significantly affects the response of the paint and should be given the uppermost attention in the case of hypersonic flows. Also, it was found that past a certain temperature threshold, the paint intensity degradation became irreversible. The comparison between the pressure tap measurement and the pressure sensitive paint showed the right trend. However, there exists a shift when it comes to the actual value. Therefore, further investigation is under way to find the cause of the shift.

Aerodynamics of an airfoil with a jet issuing from its surface

NASA Technical Reports Server (NTRS)

Tavella, D. A.; Karamcheti, K.

1982-01-01

A simple, two dimensional, incompressible and inviscid model for the problem posed by a two dimensional wing with a jet issuing from its lower surface is considered and a parametric analysis is carried out to observe how the aerodynamic characteristics depend on the different parameters. The mathematical problem constitutes a boundary value problem where the position of part of the boundary is not known a priori. A nonlinear optimization approach was used to solve the problem, and the analysis reveals interesting characteristics that may help to better understand the physics involved in more complex situations in connection with high lift systems.

Multiobjective Aerodynamic Shape Optimization Using Pareto Differential Evolution and Generalized Response Surface Metamodels

NASA Technical Reports Server (NTRS)

Madavan, Nateri K.

2004-01-01

Differential Evolution (DE) is a simple, fast, and robust evolutionary algorithm that has proven effective in determining the global optimum for several difficult single-objective optimization problems. The DE algorithm has been recently extended to multiobjective optimization problem by using a Pareto-based approach. In this paper, a Pareto DE algorithm is applied to multiobjective aerodynamic shape optimization problems that are characterized by computationally expensive objective function evaluations. To improve computational expensive the algorithm is coupled with generalized response surface meta-models based on artificial neural networks. Results are presented for some test optimization problems from the literature to demonstrate the capabilities of the method.

Investigation of space shuttle orbiter subsonic stability and control characteristics and determination of control surface hinge moments in the Rockwell International low speed wind tunnel (OA37)

NASA Technical Reports Server (NTRS)

Hughes, T.

1974-01-01

Experimental aerodynamic investigations were conducted on a string-mounted 0.030 scale representation of the 140A/B space shuttle orbiter in the 7.75- by 11-foot low speed wind tunnel. The primary test objectives were to establish basic longitudinal and lateral directional stability and control characteristics for the basic configuration plus control surface hinge moments. Aerodynamic force and moment data were measured in the body axis system by an internally mounted, six-component strain gage balance. Additional configurations investigated were sealed rudder hingeline gaps, sealed elevon gaps and compartmentized speedbrakes.

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

Recent developments in rotary-wing aerodynamic theory

NASA Technical Reports Server (NTRS)

Johnson, W.

1986-01-01

Current progress in the computational analysis of rotary-wing flowfields is surveyed, and some typical results are presented in graphs. Topics examined include potential theory, rotating coordinate systems, lifting-surface theory (moving singularity, fixed wing, and rotary wing), panel methods (surface singularity representations, integral equations, and compressible flows), transonic theory (the small-disturbance equation), wake analysis (hovering rotor-wake models and transonic blade-vortex interaction), limitations on computational aerodynamics, and viscous-flow methods (dynamic-stall theories and lifting-line theory). It is suggested that the present algorithms and advanced computers make it possible to begin working toward the ultimate goal of turbulent Navier-Stokes calculations for an entire rotorcraft.

Theoretical aerodynamic characteristics of a family of slender wing-tail-body combinations

NASA Technical Reports Server (NTRS)

Lomax, Harvard; Byrd, Paul F

1951-01-01

The aerodynamic characteristics of an airplane configuration composed of a swept-back, nearly constant chord wing and a triangular tail mounted on a cylindrical body are presented. The analysis is based on the assumption that the free-stream Mach number is near unity or that the configuration is slender. The calculations for the tail are made on the assumption that the vortex system trailing back from the wing is either a sheet lying entirely in the plane of the flat tail surface or has completely "rolled up" into two point vortices that lie either in, above, or below the plane of the tail surface.

Derivation of Nonlinear Wave Equation for Flexural Motions of AN Elastic Beam Travelling in AN Air-Filled Tube

NASA Astrophysics Data System (ADS)

Sugimoto, N.; Kugo, K.; Watanabe, Y.

2002-07-01

Asymptotic analysis is carried out to derive a nonlinear wave equation for flexural motions of an elastic beam of circular cross-section travelling along the centre-axis of an air-filled, circular tube placed coaxially. Both the beam and tube are assumed to be long enough for end-effects to be ignored and the aerodynamic loading on the lateral surface of the beam is considered. Assuming a compressible inviscid fluid, the velocity potential of the air is sought systematically in the form of power series in terms of the ratios of the tube radius to a wavelength and of a typical deflection to the radius. Evaluating the pressure force acting on the lateral surface of the beam, the aerodynamic loading including the effects of finite deflection as well as of air's compressibility and axial curvature of the beam are obtained. Although the nonlinearity arises from the kinematical condition on the beam surface, it may be attributed to the presence of the tube wall. With the aerodynamic loading thus obtained, a nonlinear wave equation is derived, whereas linear theory is assumed for the flexural motions of the beam. Some discussions are given on the results.

Improvements to the missile aerodynamic prediction code DEMON3

NASA Technical Reports Server (NTRS)

Dillenius, Marnix F. E.; Johnson, David L.; Lesieutre, Daniel J.

1992-01-01

The computer program DEMON3 was developed for the aerodynamic analysis of nonconventional supersonic configurations comprising a body with noncircular cross section and up to two wing or fin sections. Within a wing or fin section, the lifting surfaces may be cruciform, triform, planar, or low profile layouts; the planforms of the lifting surfaces allow for breaks in sweep. The body and fin sections are modeled by triplet and constant u-velocity panels, respectively, accounting for mutual body-fin interference. Fin thickness effects are included for the use of supersonic planar source panels. One of the unique features of DEMON3 is the modeling of high angle of attack vortical effects associated with the lifting surfaces and the body. In addition, shock expansion and Newtonian pressure calculation methods can be optionally engaged. These two dimensional nonlinear methods are augmented by aerodynamic interference determined from the linear panel methods. Depending on the geometric details of the body, the DEMON3 program can be used to analyze nonconventional configurations at angles of attack up to 25 degrees for Mach numbers from 1.1 to 6. Calculative results and comparisons with experimental data demonstrate the capabilities of DEMON3. Limitations and deficiencies are listed.

The aerodynamics of propellers and rotors using an acoustic formulation in the time domain

NASA Technical Reports Server (NTRS)

Long, L. N.

1983-01-01

The aerodynamics of propellers and rotors is especially complicated because of the highly three-dimensional and compressible nature of the flow field. However, in linearized theory the problem is governed by the wave equation, and a numerically-efficient integral formulation can be derived. This reduces the problem from one in space to one over a surface. Many such formulations exist in the aeroacoustics literature, but these become singular integral equations if one naively tries to use them to predict surface pressures, i.e., for aerodynamics. The present paper illustrates how one must interpret these equations in order to obtain nonambiguous results. After the regularized form of the integral equation is derived, a method for solving it numerically is described. This preliminary computer code uses Legendre-Gaussian quadrature to solve the equation. Numerical results are compared to experimental results for ellipsoids, wings, and rotors, including effects due to lift. Compressibility and the farfield boundary conditions are satisfied automatically using this method.

An integrated CFD/experimental analysis of aerodynamic forces and moments

NASA Technical Reports Server (NTRS)

Melton, John E.; Robertson, David D.; Moyer, Seth A.

1989-01-01

Aerodynamic analysis using computational fluid dynamics (CFD) is most fruitful when it is combined with a thorough program of wind tunnel testing. The understanding of aerodynamic phenomena is enhanced by the synergistic use of both analysis methods. A technique is described for an integrated approach to determining the forces and moments acting on a wind tunnel model by using a combination of experimentally measured pressures and CFD predictions. The CFD code used was FLO57 (an Euler solver) and the wind tunnel model was a heavily instrumented delta wing with 62.5 deg of leading-edge sweep. A thorough comparison of the CFD results and the experimental data is presented for surface pressure distributions and longitudinal forces and moments. The experimental pressures were also integrated over the surface of the model and the resulting forces and moments are compared to the CFD and wind tunnel results. The accurate determination of various drag increments via the combined use of the CFD and experimental pressures is presented in detail.

Design, develop and test high temperature dynamic seals for the space shuttle's aerodynamic control surfaces

NASA Technical Reports Server (NTRS)

1973-01-01

A description is given of the design, development and testing of high temperature dynamic seals for the gaps between the structure and aerodynamic control surfaces on the space shuttle. These aerodynamic seals are required to prevent high temperature airflow from damaging thermally unprotected structures and components during entry. Two seal concepts evolved a curtain seal for the spanwise elevon cove gap, and a labyrinth seal for the area above the elevon, at the gap between the end of the elevon and the fuselage. On the basis of development testing, both seal concepts were shown to be feasible for controlling internal temperatures to 350 F or less when exposed to a typical space shuttle entry environment. The curtain seal concept demonstrated excellent test results and merits strong consideration for application on the space shuttle orbiter. The labyrinth seal concept, although demonstrating significant temperature reduction characteristics, may or may not be required on the Orbiter, depending on the actual design configuration and flight environment.

Optimized aerodynamic design process for subsonic transport wing fitted with winglets. [wind tunnel model

NASA Technical Reports Server (NTRS)

Kuhlman, J. M.

1979-01-01

The aerodynamic design of a wind-tunnel model of a wing representative of that of a subsonic jet transport aircraft, fitted with winglets, was performed using two recently developed optimal wing-design computer programs. Both potential flow codes use a vortex lattice representation of the near-field of the aerodynamic surfaces for determination of the required mean camber surfaces for minimum induced drag, and both codes use far-field induced drag minimization procedures to obtain the required spanloads. One code uses a discrete vortex wake model for this far-field drag computation, while the second uses a 2-D advanced panel wake model. Wing camber shapes for the two codes are very similar, but the resulting winglet camber shapes differ widely. Design techniques and considerations for these two wind-tunnel models are detailed, including a description of the necessary modifications of the design geometry to format it for use by a numerically controlled machine for the actual model construction.

The Aerodynamic Plane Table

NASA Technical Reports Server (NTRS)

Zahm, A F

1924-01-01

This report gives the description and the use of a specially designed aerodynamic plane table. For the accurate and expeditious geometrical measurement of models in an aerodynamic laboratory, and for miscellaneous truing operations, there is frequent need for a specially equipped plan table. For example, one may have to measure truly to 0.001 inch the offsets of an airfoil at many parts of its surface. Or the offsets of a strut, airship hull, or other carefully formed figure may require exact calipering. Again, a complete airplane model may have to be adjusted for correct incidence at all parts of its surfaces or verified in those parts for conformance to specifications. Such work, if but occasional, may be done on a planing or milling machine; but if frequent, justifies the provision of a special table. For this reason it was found desirable in 1918 to make the table described in this report and to equip it with such gauges and measures as the work should require.

An Efficient Radial Basis Function Mesh Deformation Scheme within an Adjoint-Based Aerodynamic Optimization Framework

NASA Astrophysics Data System (ADS)

Poirier, Vincent

Mesh deformation schemes play an important role in numerical aerodynamic optimization. As the aerodynamic shape changes, the computational mesh must adapt to conform to the deformed geometry. In this work, an extension to an existing fast and robust Radial Basis Function (RBF) mesh movement scheme is presented. Using a reduced set of surface points to define the mesh deformation increases the efficiency of the RBF method; however, at the cost of introducing errors into the parameterization by not recovering the exact displacement of all surface points. A secondary mesh movement is implemented, within an adjoint-based optimization framework, to eliminate these errors. The proposed scheme is tested within a 3D Euler flow by reducing the pressure drag while maintaining lift of a wing-body configured Boeing-747 and an Onera-M6 wing. As well, an inverse pressure design is executed on the Onera-M6 wing and an inverse span loading case is presented for a wing-body configured DLR-F6 aircraft.

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.

Metal-wool heat shields for space shuttle. [design, fabrication, and attachment to structure

NASA Technical Reports Server (NTRS)

Miller, R. C.; Clure, J. L.

1974-01-01

The packaging of metal wool for reusable thermal heat shields applied to aerodynamic and other surfaces for the space shuttle was analyzed and designed, and samples were fabricated and experimentally studied. Parametric trends were prepared for selected configurations. An all-metal thermally efficient, reliable, reusable and producible heat shield system was designed and structurally tested for use on spacecraft aerodynamic surfaces where temperatures do not exceed 810 K. Stainless steel sheet, primarily for structure and secondarily in the transverse plane for thermal expansion, was shown to accommodate thermal expansion in all directions when restrained at the edges and heated to 1360 K. Aerodynamic loads of 0.35 x 1000,000 newtons/sq meter, and higher, may be easily accepted by structures of this design. Seven all-metal thermal protection specimens, 12.7 cm square and 2.5 cm thick were fabricated and are being experimentally evaluated at simulated shuttle entry conditions in an arc jet facility.

Combined experiment Phase 2 data characterization

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

Miller, M.S.; Shipley, D.E.; Young, T.S.

1995-11-01

The National Renewable Energy Laboratory`s ``Combined Experiment`` has yielded a large quantity of experimental data on the operation of a downwind horizontal axis wind turbine under field conditions. To fully utilize this valuable resource and identify particular episodes of interest, a number of databases were created that characterize individual data events and rotational cycles over a wide range of parameters. Each of the 59 five-minute data episodes collected during Phase 11 of the Combined Experiment have been characterized by the mean, minimum, maximum, and standard deviation of all data channels, except the blade surface pressures. Inflow condition, aerodynamic force coefficient,more » and minimum leading edge pressure coefficient databases have also been established, characterizing each of nearly 21,000 blade rotational cycles. In addition, a number of tools have been developed for searching these databases for particular episodes of interest. Due to their extensive size, only a portion of the episode characterization databases are included in an appendix, and examples of the cycle characterization databases are given. The search tools are discussed and the FORTRAN or C code for each is included in appendices.« less

Analytic corrections to CFD heating predictions accounting for changes in surface catalysis

NASA Technical Reports Server (NTRS)

Gnoffo, Peter A.; Inger, George R.

1996-01-01

Integral boundary-layer solution techniques applicable to the problem of determining aerodynamic heating rates of hypersonic vehicles in the vicinity of stagnation points and windward centerlines are briefly summarized. A new approach for combining the insight afforded by integral boundary-layer analysis with comprehensive (but time intensive) computational fluid dynamic (CFD) flowfield solutions of the thin-layer Navier-Stokes equations is described. The approach extracts CFD derived quantities at the wall and at the boundary layer edge for inclusion in a post-processing boundary-layer analysis. It allows a designer at a workstation to address two questions, given a single CFD solution. (1) How much does the heating change for a thermal protection system with different catalytic properties than was used in the original CFD solution? (2) How does the heating change at the interface of two different TPS materials with an abrupt change in catalytic efficiency? The answer to the second question is particularly important, because abrupt changes from low to high catalytic efficiency can lead to localized increase in heating which exceeds the usually conservative estimate provided by a fully catalytic wall assumption.

Supersonic Aerodynamic Characteristics of Blunt Body Trim Tab Configurations

NASA Technical Reports Server (NTRS)

Korzun, Ashley M.; Murphy, Kelly J.; Edquist, Karl T.

2013-01-01

Trim tabs are aerodynamic control surfaces that can allow an entry vehicle to meet aerodynamic performance requirements while reducing or eliminating the use of ballast mass and providing a capability to modulate the lift-to-drag ratio during entry. Force and moment data were obtained on 38 unique, blunt body trim tab configurations in the NASA Langley Research Center Unitary Plan Wind Tunnel. The data were used to parametrically assess the supersonic aerodynamic performance of trim tabs and to understand the influence of tab area, cant angle, and aspect ratio. Across the range of conditions tested (Mach numbers of 2.5, 3.5, and 4.5; angles of attack from -4deg to +20deg; angles of sideslip from 0deg to +8deg), the effects of varying tab area and tab cant angle were found to be much more significant than effects from varying tab aspect ratio. Aerodynamic characteristics exhibited variation with Mach number and forebody geometry over the range of conditions tested. Overall, the results demonstrate that trim tabs are a viable approach to satisfy aerodynamic performance requirements of blunt body entry vehicles with minimal ballast mass. For a 70deg sphere-cone, a tab with 3% area of the forebody and canted approximately 35deg with no ballast mass was found to give the same trim aerodynamics as a baseline model with ballast mass that was 5% of the total entry mass.

Rotational Augmentation on a 2.3 MW Rotor Blade with Thick Flatback Airfoil Cross-Sections: Preprint

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

Schreck, S.; Fingersh, L.; Siegel, K.

2013-01-01

Rotational augmentation was analyzed for a 2.3 MW wind turbine, which was equipped with thick flatback airfoils at inboard radial locations and extensively instrumented for acquisition of time varying surface pressures. Mean aerodynamic force and surface pressure data were extracted from an extensive field test database, subject to stringent criteria for wind inflow and turbine operating conditions. Analyses of these data showed pronounced amplification of aerodynamic forces and significant enhancements to surface pressures in response to rotational influences, relative to two-dimensional, stationary conditions. Rotational augmentation occurrence and intensity in the current effort was found to be consistent with that observedmore » in previous research. Notably, elevated airfoil thickness and flatback design did not impede rotational augmentation.« less

Effects of structural nonlinearity on subsonic aeroelastic characteristics of an aircraft wing with control surface

NASA Astrophysics Data System (ADS)

Bae, J.-S.; Inman, D. J.; Lee, I.

2004-07-01

The nonlinear aeroelastic characteristics of an aircraft wing with a control surface are investigated. A doublet-hybrid method is used for the calculation of subsonic unsteady aerodynamic forces and the minimum-state approximation is used for the approximation of aerodynamic forces. A free vibration analysis is performed using the finite element and the fictitious mass methods. The structural nonlinearity in the control surface hinge is represented by both free-play and a bilinear nonlinearity. These nonlinearities are linearized using the describing function method. From the nonlinear flutter analysis, various types of limit cycle oscillations and periodic motions are observed in a wide range of air speeds below the linear flutter boundary. The effects of structural nonlinearities on aeroelastic characteristics are investigated.

A new aerodynamic integral equation based on an acoustic formula in the time domain

NASA Technical Reports Server (NTRS)

Farassat, F.

1984-01-01

An aerodynamic integral equation for bodies moving at transonic and supersonic speeds is presented. Based on a time-dependent acoustic formula for calculating the noise emanating from the outer portion of a propeller blade travelling at high speed (the Ffowcs Williams-Hawking formulation), the loading terms and a conventional thickness source terms are retained. Two surface and three line integrals are employed to solve an equation for the loading noise. The near-field term is regularized using the collapsing sphere approach to obtain semiconvergence on the blade surface. A singular integral equation is thereby derived for the unknown surface pressure, and is amenable to numerical solutions using Galerkin or collocation methods. The technique is useful for studying the nonuniform inflow to the propeller.

Experimental Hypersonic Aerodynamic Characteristics of the Space Shuttle Orbiter for a Range of Damage Scenarios

NASA Technical Reports Server (NTRS)

Brauckman, Gregory J.; Scallion, William I.

2003-01-01

Aerodynamic tests in support of the Columbia accident investigation were conducted in two hypersonic wind tunnels at the NASA Langley Research Center, the 20-Inch Mach 6 Air Tunnel and the 20-Inch Mach 6 CF4 Tunnel. The primary purpose of these tests was to measure the forces and moments generated by a variety of outer mold line alterations (damage scenarios) using 0.0075-scale models of the Space Shuttle Orbiter (approximately 10 inches in length). Simultaneously acquired global heat transfer mappings were obtained for a majority of the configurations tested. Test parameters include angles of attack from 38 to 42 deg, unit Reynolds numbers from 0.26 to 3.0 x10^6 per foot, and normal shock density ratios of 5 (Mach 6 air) and 12 (Mach 6 CF4). The damage scenarios evaluated included asymmetric boundary layer transition, gouges in the windward surface acreage thermal protection system tiles, wing leading edge damage (partially and fully missing reinforced carbon-carbon (RCC) panels), holes through the wing from the windward surface to the leeside, deformation of the wing windward surface, and main landing gear door and/or gear deployment. The aerodynamic data were compared to the magnitudes and directions observed in flight, and the heating images were evaluated in terms of the location of the generated disturbances and how these disturbance might relate to the response of discrete gages on the Columbia Orbiter vehicle during entry. The measured aerodynamic increments were generally small in magnitude, as were the flight-derived values during most of the entry. Asymmetric boundary layer transition (ABLT) results were consistent with the flight-derived Shuttle ABLT model, but not with the observed flight trends for STS-107. The partially missing leading edge panel results best matched both the early aerodynamic and heating trends observed in flight. A progressive damage scenario is presented that qualitatively matches the flight observations for the full entry.

Assessment of CFD-based Response Surface Model for Ares I Supersonic Ascent Aerodynamics

NASA Technical Reports Server (NTRS)

Hanke, Jeremy L.

2011-01-01

The Ascent Force and Moment Aerodynamic (AFMA) Databases (DBs) for the Ares I Crew Launch Vehicle (CLV) were typically based on wind tunnel (WT) data, with increments provided by computational fluid dynamics (CFD) simulations for aspects of the vehicle that could not be tested in the WT tests. During the Design Analysis Cycle 3 analysis for the outer mold line (OML) geometry designated A106, a major tunnel mishap delayed the WT test for supersonic Mach numbers (M) greater than 1.6 in the Unitary Plan Wind Tunnel at NASA Langley Research Center, and the test delay pushed the final delivery of the A106 AFMA DB back by several months. The aero team developed an interim database based entirely on the already completed CFD simulations to mitigate the impact of the delay. This CFD-based database used a response surface methodology based on radial basis functions to predict the aerodynamic coefficients for M > 1.6 based on only the CFD data from both WT and flight Reynolds number conditions. The aero team used extensive knowledge of the previous AFMA DB for the A103 OML to guide the development of the CFD-based A106 AFMA DB. This report details the development of the CFD-based A106 Supersonic AFMA DB, constructs a prediction of the database uncertainty using data available at the time of development, and assesses the overall quality of the CFD-based DB both qualitatively and quantitatively. This assessment confirms that a reasonable aerodynamic database can be constructed for launch vehicles at supersonic conditions using only CFD data if sufficient knowledge of the physics and expected behavior is available. This report also demonstrates the applicability of non-parametric response surface modeling using radial basis functions for development of aerodynamic databases that exhibit both linear and non-linear behavior throughout a large data space.

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.

Advanced Aerodynamic Design of Passive Porosity Control Effectors

NASA Technical Reports Server (NTRS)

Hunter, Craig A.; Viken, Sally A.; Wood, Richard M.; Bauer, Steven X. S.

2001-01-01

This paper describes aerodynamic design work aimed at developing a passive porosity control effector system for a generic tailless fighter aircraft. As part of this work, a computational design tool was developed and used to layout passive porosity effector systems for longitudinal and lateral-directional control at a low-speed, high angle of attack condition. Aerodynamic analysis was conducted using the NASA Langley computational fluid dynamics code USM3D, in conjunction with a newly formulated surface boundary condition for passive porosity. Results indicate that passive porosity effectors can provide maneuver control increments that equal and exceed those of conventional aerodynamic effectors for low-speed, high-alpha flight, with control levels that are a linear function of porous area. This work demonstrates the tremendous potential of passive porosity to yield simple control effector systems that have no external moving parts and will preserve an aircraft's fixed outer mold line.

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.

Building Aerodynamic Databases for the SLS Design Process

NASA Technical Reports Server (NTRS)

Rogers, Stuart; Dalle, Derek J.; Lee, Henry; Meeroff, Jamie; Onufer, Jeffrey; Chan, William; Pulliam, Thomas

2017-01-01

NASA's new Space Launch System (SLS) will be the first rocket since the Saturn V (1967-1973) to carry astronauts beyond low earth orbit-and will carry 10% more payload than Saturn V and three times the payload of the space shuttle. The SLS configuration consists of a center core and two solid rocket boosters that separate from the core as their fuel is exhausted two minutes after lift-off. During these first two minutes of flight, the vehicle powers its way through strong shock waves as it accelerates past the speed of sound, then pushes beyond strong aerodynamic loads at the maximum dynamic pressure, and is ultimately enveloped by gaseous plumes from the booster-separation motors. The SLS program relies on computational fluid dynamic (CFD) simulations to provide much of the data needed to build aerodynamic databases describing the structural load distribution, surface pressures, and aerodynamic forces on the vehicle.

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.

Aerodynamic Shape Optimization Design of Wing-Body Configuration Using a Hybrid FFD-RBF Parameterization Approach

NASA Astrophysics Data System (ADS)

Liu, Yuefeng; Duan, Zhuoyi; Chen, Song

2017-10-01

Aerodynamic shape optimization design aiming at improving the efficiency of an aircraft has always been a challenging task, especially when the configuration is complex. In this paper, a hybrid FFD-RBF surface parameterization approach has been proposed for designing a civil transport wing-body configuration. This approach is simple and efficient, with the FFD technique used for parameterizing the wing shape and the RBF interpolation approach used for handling the wing body junction part updating. Furthermore, combined with Cuckoo Search algorithm and Kriging surrogate model with expected improvement adaptive sampling criterion, an aerodynamic shape optimization design system has been established. Finally, the aerodynamic shape optimization design on DLR F4 wing-body configuration has been carried out as a study case, and the result has shown that the approach proposed in this paper is of good effectiveness.

Europa Kinetic Ice Penetrator System for Hyper Velocity Instrument Deposition

NASA Astrophysics Data System (ADS)

Robinson, Tessa

Landing of a payload on any celestial body has only used a soft landing system. These systems use retro rockets and atmospheric components to match velocity and then overcome local gravity in order to land on the surface. This is a proposed system for depositing instrumentation on an icy surface at hypervelocity using the properties of different projectiles and ejecta properties that would negate the need for a soft landing system. This system uses two projectiles, a cylinder with inner aerodynamic surfaces and a payload section with a conical nose and aerodynamic surfaces. The cylinder lands first, creates a region of fractured ice, and directs that fractured material into a collimated ejecta stream. The payload travels through the ejecta and lands in the fractured region. The combination of the ejecta stream and the softened target material reduces the impact acceleration to within survivable levels.

Effect of the Surface Condition of a Wing on the Aerodynamic Characteristics of an Airplane

NASA Technical Reports Server (NTRS)

Defrance, S J

1934-01-01

In order to determine the effect of the surface conditions of a wing on the aerodynamic characteristics of an airplane, tests were conducted in the N.A.C.A. full-scale wind tunnel on the Fairchild F-22 airplane first with normal commercial finish of wing surface and later with the same wing polished. Comparison of the characteristics of the airplane with the two surface conditions shows that the polish caused a negligible change in the lift curve, but reduced the minimum drag coefficient by 0.001. This reduction in drag if applied to an airplane with a given speed of 200 miles per hour and a minimum drag coefficient of 0.025 would increase the speed only 2.9 miles per hour, but if the speed remained the same, the power would be reduced 4 percent.

Effect of control surface mass unbalance on the stability of a closed-loop active control system

NASA Technical Reports Server (NTRS)

Nissim, E.

1989-01-01

The effects on stability of inertial forces arising from closed-loop activation of mass-unbalanced control surfaces are studied analytically using inertial energy approach, similar to the aerodynamic energy approach used for flutter suppression. The limitations of a single control surface like a leading-edge (LE) control or a trailing-edge (TE) control are demonstrated and compared to the superior combined LE-TE mass unbalanced system. It is shown that a spanwise section for sensor location can be determined which ensures minimum sensitivity to the mode shapes of the aircraft. It is shown that an LE control exhibits compatibility between inertial stabilization and aerodynamic stabilization, and that a TE control lacks such compatibility. The results of the present work should prove valuable, both for the purpose of flutter suppression using mass unbalanced control surfaces, or for the stabilization of structural modes of large space structures by means of inertial forces.

Aerodynamic Design Using Neural Networks

NASA Technical Reports Server (NTRS)

Rai, Man Mohan; Madavan, Nateri K.

2003-01-01

The design of aerodynamic components of aircraft, such as wings or engines, involves a process of obtaining the most optimal component shape that can deliver the desired level of component performance, subject to various constraints, e.g., total weight or cost, that the component must satisfy. Aerodynamic design can thus be formulated as an optimization problem that involves the minimization of an objective function subject to constraints. A new aerodynamic design optimization procedure based on neural networks and response surface methodology (RSM) incorporates the advantages of both traditional RSM and neural networks. The procedure uses a strategy, denoted parameter-based partitioning of the design space, to construct a sequence of response surfaces based on both neural networks and polynomial fits to traverse the design space in search of the optimal solution. Some desirable characteristics of the new design optimization procedure include the ability to handle a variety of design objectives, easily impose constraints, and incorporate design guidelines and rules of thumb. It provides an infrastructure for variable fidelity analysis and reduces the cost of computation by using less-expensive, lower fidelity simulations in the early stages of the design evolution. The initial or starting design can be far from optimal. The procedure is easy and economical to use in large-dimensional design space and can be used to perform design tradeoff studies rapidly. Designs involving multiple disciplines can also be optimized. Some practical applications of the design procedure that have demonstrated some of its capabilities include the inverse design of an optimal turbine airfoil starting from a generic shape and the redesign of transonic turbines to improve their unsteady aerodynamic characteristics.

Aerodynamic losses calculation of a turbine blade with film cooling with forward and backward injection by numerical method

NASA Astrophysics Data System (ADS)

Prajapati, Anil

Thermal efficiency and power output of gas turbines can be increased by increasing the turbine blade inlet temperature. However, the main problem is the durability of the turbine blade due to the thermal stress on it at high temperature. This has led to the development of film cooling technology, in which coolant is injected from a series of cooling holes made on the blade surface to form an insulating blanket over the blade surface. However, it has to pay the aerodynamic penalties due to the injection of coolant, which are not fully understood. Pressure loss coefficient is one of the easy and widely used parameters to determine the aerodynamic loss occurred on a turbine blade. The losses occurred on the turbine blade with forward injection and backward injection cooling are studied at a different blowing ratios by a numerical simulation, which shows that the loss is higher in the case of backward injection than in forward injection. Fan-shaped cooling holes are also considered to compare with the cylindrical holes. It is observed that the loss is increased due to the fan-shaped holes in the forward injection whereas there is not a substantial difference due to the fan-shaped holes in the backward injection. The aerodynamic loss due to the location of coolant injection is studied by using injection from the leading edge, pressure side, suction side and trailing edge respectively. The study is performed to determine the effect of incidence angles and coolant injection angles on the aerodynamic loss.

Dustiness of Fine and Nanoscale Powders

PubMed Central

Evans, Douglas E.; Baron, Paul A.

2013-01-01

Dustiness may be defined as the propensity of a powder to form airborne dust by a prescribed mechanical stimulus; dustiness testing is typically intended to replicate mechanisms of dust generation encountered in workplaces. A novel dustiness testing device, developed for pharmaceutical application, was evaluated in the dustiness investigation of 27 fine and nanoscale powders. The device efficiently dispersed small (mg) quantities of a wide variety of fine and nanoscale powders, into a small sampling chamber. Measurements consisted of gravimetrically determined total and respirable dustiness. The following materials were studied: single and multiwalled carbon nanotubes, carbon nanofibers, and carbon blacks; fumed oxides of titanium, aluminum, silicon, and cerium; metallic nanoparticles (nickel, cobalt, manganese, and silver) silicon carbide, Arizona road dust; nanoclays; and lithium titanate. Both the total and respirable dustiness spanned two orders of magnitude (0.3–37.9% and 0.1–31.8% of the predispersed test powders, respectively). For many powders, a significant respirable dustiness was observed. For most powders studied, the respirable dustiness accounted for approximately one-third of the total dustiness. It is believed that this relationship holds for many fine and nanoscale test powders (i.e. those primarily selected for this study), but may not hold for coarse powders. Neither total nor respirable dustiness was found to be correlated with BET surface area, therefore dustiness is not determined by primary particle size. For a subset of test powders, aerodynamic particle size distributions by number were measured (with an electrical low-pressure impactor and an aerodynamic particle sizer). Particle size modes ranged from approximately 300nm to several micrometers, but no modes below 100nm, were observed. It is therefore unlikely that these materials would exhibit a substantial sub-100nm particle contribution in a workplace. PMID:23065675

Mars 2020 Entry, Descent and Landing Instrumentation (MEDLI2)

NASA Technical Reports Server (NTRS)

Bose, Deepak; Wright, Henry; White, Todd; Schoenenberger, Mark; Santos, Jose; Karlgaard, Chris; Kuhl, Chris; Oishi, TOmo; Trombetta, Dominic

2016-01-01

This paper will introduce Mars Entry Descent and Landing Instrumentation (MEDLI2) on NASA's Mars2020 mission. Mars2020 is a flagship NASA mission with science and technology objectives to help answer questions about possibility of life on Mars as well as to demonstrate technologies for future human expedition. Mars2020 is scheduled for launch in 2020. MEDLI2 is a suite of instruments embedded in the heatshield and backshell thermal protection systems of Mars2020 entry vehicle. The objectives of MEDLI2 are to gather critical aerodynamics, aerothermodynamics and TPS performance data during EDL phase of the mission. MEDLI2 builds up the success of MEDLI flight instrumentation on Mars Science Laboratory mission in 2012. MEDLI instrumentation suite measured surface pressure and TPS temperature on the heatshield during MSL entry into Mars. MEDLI data has since been used for unprecedented reconstruction of aerodynamic drag, vehicle attitude, in-situ atmospheric density, aerothermal heating, transition to turbulence, in-depth TPS performance and TPS ablation. [1,2] In addition to validating predictive models, MEDLI data has highlighted extra margin available in the MSL forebody TPS, which can potentially be used to reduce vehicle parasitic mass. MEDLI2 expands the scope of instrumentation by focusing on quantities of interest not addressed in MEDLI suite. The type the sensors are expanded and their layout on the TPS modified to meet these new objectives. The paper will provide key motivation and governing requirements that drive the choice and the implementation of the new sensor suite. The implementation considerations of sensor selection, qualification, and demonstration of minimal risk to the host mission will be described. The additional challenges associated with mechanical accommodation, electrical impact, data storage and retrieval for MEDLI2 system, which extends sensors to backshell will also be described.

Aerodynamic characteristics of some lifting reentry concepts applicable to transatmospheric vehicle design studies

NASA Technical Reports Server (NTRS)

Spearman, M. L.

1984-01-01

The aerodynamic characteristics of some lifting reentry concepts are examined with a view to the applicability of such concepts to the design of possible transatmospheric vehicles (TAV). A considerable amount of research has been done in past years with vehicle concepts suitable for manned atmospheric-entry, atmospheric flight, and landing. Some of the features of these concepts that permit flight in or out of the atmosphere with maneuver capability should be useful in the mission requirements of TAV's. The concepts illustrated include some hypersonic-body shapes with and without variable geometry surfaces, and a blunt lifting-body configuration. The merits of these concepts relative to the aerodynamic behavior of a TAV are discussed.

Effect of underwing aft-mounted nacelles on the longitudinal aerodynamic characteristics of a high-wing transport airplane

NASA Technical Reports Server (NTRS)

Abeyounis, W. K.; Patterson, J. C., Jr.

1985-01-01

As part of a propulsion/airframe integration program, tests were conducted in the Langley 16-Foot Transonic Tunnel to determine the longitudinal aerodynamic effects of installing flow through engine nacelles in the aft underwing position of a high wing transonic transfer airplane. Mixed flow nacelles with circular and D-shaped inlets were tested at free stream Mach numbers from 0.70 to 0.85 and angles of attack from -2.5 deg to 4.0 deg. The aerodynamic effects of installing antishock bodies on the wing and nacelle upper surfaces as a means of attaching and supporting nacelles in an extreme aft position were investigated.

Preliminary Studies on Aerodynamic Control with Direct Current Discharge at Hypersonic Speed

NASA Astrophysics Data System (ADS)

Watanabe, Yasumasa; Takama, Yoshiki; Imamura, Osamu; Watanuki, Tadaharu; Suzuki, Kojiro

A new idea of an aerodynamic control device for hypersonic vehicles using plasma discharges is presented. The effect of DC plasma discharge on a hypersonic flow is examined with both experiments and CFD analyses. It is revealed that the surface pressure upstream of plasma area significantly increases, which would be preferable in realizing a new aerodynamic control devices. Such pressure rise is also observed in the result of analyses of the Navier-Stokes equations with energy addition that simulates the Joule heating of a plasma discharge. It is revealed that the pressure rise due to the existence of the plasma discharge can be qualitatively explained as an effect of Joule heating.

Sensitivity analysis of a wing aeroelastic response

NASA Technical Reports Server (NTRS)

Kapania, Rakesh K.; Eldred, Lloyd B.; Barthelemy, Jean-Francois M.

1991-01-01

A variation of Sobieski's Global Sensitivity Equations (GSE) approach is implemented to obtain the sensitivity of the static aeroelastic response of a three-dimensional wing model. The formulation is quite general and accepts any aerodynamics and structural analysis capability. An interface code is written to convert one analysis's output to the other's input, and visa versa. Local sensitivity derivatives are calculated by either analytic methods or finite difference techniques. A program to combine the local sensitivities, such as the sensitivity of the stiffness matrix or the aerodynamic kernel matrix, into global sensitivity derivatives is developed. The aerodynamic analysis package FAST, using a lifting surface theory, and a structural package, ELAPS, implementing Giles' equivalent plate model are used.

Computational Aerodynamics of Shuttle Orbiter Damage Scenarios in Support of the Columbia Accident Investigation

NASA Technical Reports Server (NTRS)

Bibb, Karen L.; Prabhu, Ramadas K.

2004-01-01

In support of the Columbia Accident Investigation, inviscid computations of the aerodynamic characteristics for various Shuttle Orbiter damage scenarios were performed using the FELISA unstructured CFD solver. Computed delta aerodynamics were compared with the reconstructed delta aerodynamics in order to postulate a progression of damage through the flight trajectory. By performing computations at hypervelocity flight and CF4 tunnel conditions, a bridge was provided between wind tunnel testing in Langley's 20-Inch CF4 facility and the flight environment experienced by Columbia during re-entry. The rapid modeling capability of the unstructured methodology allowed the computational effort to keep pace with the wind tunnel and, at times, guide the wind tunnel efforts. These computations provided a detailed view of the flowfield characteristics and the contribution of orbiter components (such as the vertical tail and wing) to aerodynamic forces and moments that were unavailable from wind tunnel testing. The damage scenarios are grouped into three categories. Initially, single and multiple missing full RCC panels were analyzed to determine the effect of damage location and magnitude on the aerodynamics. Next is a series of cases with progressive damage, increasing in severity, in the region of RCC panel 9. The final group is a set of wing leading edge and windward surface deformations that model possible structural deformation of the wing skin due to internal heating of the wing structure. By matching the aerodynamics from selected damage scenarios to the reconstructed flight aerodynamics, a progression of damage that is consistent with the flight data, debris forensics, and wind tunnel data is postulated.

Laser Scanning System for Pressure and Temperature Paints

NASA Technical Reports Server (NTRS)

Sullivan, John

1997-01-01

Acquiring pressure maps of aerodynamic surfaces is very important for improving and validating the performance of aerospace vehicles. Traditional pressure measurements are taken with pressure taps embedded in the model surface that are connected to transducers. While pressure taps allow highly accurate measurements to be acquired, they do have several drawbacks. Pressure taps do not give good spatial resolution due to the need for individual pressure tubes, compounded by limited space available inside models. Also, building a model proves very costly if taps are needed because of the large amount of labor necessary to drill, connect and test each one. The typical cost to install one tap is about $200. Recently, a new method for measuring pressure on aerodynamic surfaces has been developed utilizing a technology known as pressure sensitive paints (PSP). Using PSP, pressure distributions can be acquired optically with high spatial resolution and simple model preparation. Flow structures can be easily visualized using PSP, but are missed using low spatial resolution arrays of pressure taps. PSP even allows pressure distributions to be found on rotating machinery where previously this has been extremely difficult or even impossible. The goal of this research is to develop a laser scanning system for use with pressure sensitive paints that allows accurate pressure measurements to be obtained on various aerodynamic surfaces ranging from wind tunnel models to high speed jet engine compressor blades.

Navier-Stokes Computations of a Wing-Flap Model With Blowing Normal to the Flap Surface

NASA Technical Reports Server (NTRS)

Boyd, D. Douglas, Jr.

2005-01-01

A computational study of a generic wing with a half span flap shows the mean flow effects of several blown flap configurations. The effort compares and contrasts the thin-layer, Reynolds averaged, Navier-Stokes solutions of a baseline wing-flap configuration with configurations that have blowing normal to the flap surface through small slits near the flap side edge. Vorticity contours reveal a dual vortex structure at the flap side edge for all cases. The dual vortex merges into a single vortex at approximately the mid-flap chord location. Upper surface blowing reduces the strength of the merged vortex and moves the vortex away from the upper edge. Lower surface blowing thickens the lower shear layer and weakens the merged vortex, but not as much as upper surface blowing. Side surface blowing forces the lower surface vortex farther outboard of the flap edge by effectively increasing the aerodynamic span of the flap. It is seen that there is no global aerodynamic penalty or benefit from the particular blowing configurations examined.

Estimation of wing nonlinear aerodynamic characteristics at supersonic speeds

NASA Technical Reports Server (NTRS)

Carlson, H. W.; Mack, R. J.

1980-01-01

A computational system for estimation of nonlinear aerodynamic characteristics of wings at supersonic speeds was developed and was incorporated in a computer program. This corrected linearized theory method accounts for nonlinearities in the variation of basic pressure loadings with local surface slopes, predicts the degree of attainment of theoretical leading edge thrust, and provides an estimate of detached leading edge vortex loadings that result when the theoretical thrust forces are not fully realized.

A computational system for aerodynamic design and analysis of supersonic aircraft. Part 1: General description and theoretical development

NASA Technical Reports Server (NTRS)

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

1976-01-01

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. Schematics of the program structure and the individual overlays and subroutines are described.

Hydrodynamic and aerodynamic breakup of liquid sheets

NASA Technical Reports Server (NTRS)

Ingebo, R.

1982-01-01

The effect of hydrodynamic, aerodynamic and liquid surface forces on the mean drop diameter of water sprays that are produced by the breakup of nonswirling and swirling water sheets in quiescent air and in airflows similar to those encountered in gas turbine combustors is investigated. The mean drop diameter is used to characterize fuel sprays and it is a very important factor in determining the performance and exhaust emissions of gas turbine combustors.

A new unified approach to analyze wing-body-tail configurations with control surfaces in steady, oscillatory and fully unsteady, subsonic and supersonic flows

NASA Technical Reports Server (NTRS)

Tseng, K.; Morino, L.

1975-01-01

A general formulation for the analysis of steady and unsteady, subsonic and supersonic potential aerodynamics for arbitrary complex geometries is presented. The theoretical formulation, the numerical procedure, and numerical results are included. In particular, generalized forces for fully unsteady (complex frequency) aerodynamics for an AGARD coplanar wing-tail interfering configuration in both subsonic and supersonic flows are considered.

Investigation of Phenomena of Discrete Wingtip Jets

DTIC Science & Technology

1988-08-01

larger than that in no-blowing case, this implied that the aerodynamic loading of the wing model increased in latter case. 3.3. SURFACE PRESSURE...results show that the improvement in the pressure distribution was different from that of the winglet . The winglet utilizes the principle of pressure...Ayers, R. F. and Wilde, M. R., " An experimental investigation of the aerodynamic characteristics of a low aspect ratio swept wing with blowing in a

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.

Aerodynamic roughness: A simple and alternative metric to detect the seasonality of canopy structure using flux-tower data

NASA Astrophysics Data System (ADS)

Chu, H.; Baldocchi, D. D.

2017-12-01

FLUXNET - the global network of eddy covariance tower sites provides valuable datasets of the direct and in situ measurements of fluxes and ancillary variables that are used across different disciplines and applications. Aerodynamic roughness (i.e., roughness length, zero plane displacement height) are one of the potential parameters that can be derived from flux-tower data and are crucial for the applications of land surface models and flux footprint models. As aerodynamic roughness are tightly associated with canopy structures (e.g., canopy height, leaf area), such parameters could potentially serve as an alternative metric for detecting the change of canopy structure (e.g., change of leaf areas in deciduous ecosystems). This study proposes a simple approach for deriving aerodynamic roughness from flux-tower data, and tests their suitability and robustness in detecting the seasonality of canopy structure. We run tests across a broad range of deciduous forests, and compare the seasonality derived from aerodynamic roughness (i.e., starting and ending dates of leaf-on period and peak-foliage period) against those obtained from remote sensing or in situ leaf area measurements. Our findings show aerodynamic roughness generally captures the timing of changes of leaf areas in deciduous forests. Yet, caution needs to be exercised while interpreting the absolute values of the roughness estimates.

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.

Touchdown to take-off: at the interface of flight and surface locomotion

PubMed Central

2017-01-01

Small aerial robots are limited to short mission times because aerodynamic and energy conversion efficiency diminish with scale. One way to extend mission times is to perch, as biological flyers do. Beyond perching, small robot flyers benefit from manoeuvring on surfaces for a diverse set of tasks, including exploration, inspection and collection of samples. These opportunities have prompted an interest in bimodal aerial and surface locomotion on both engineered and natural surfaces. To accomplish such novel robot behaviours, recent efforts have included advancing our understanding of the aerodynamics of surface approach and take-off, the contact dynamics of perching and attachment and making surface locomotion more efficient and robust. While current aerial robots show promise, flying animals, including insects, bats and birds, far surpass them in versatility, reliability and robustness. The maximal size of both perching animals and robots is limited by scaling laws for both adhesion and claw-based surface attachment. Biomechanists can use the current variety of specialized robots as inspiration for probing unknown aspects of bimodal animal locomotion. Similarly, the pitch-up landing manoeuvres and surface attachment techniques of animals can offer an evolutionary design guide for developing robots that perch on more diverse and complex surfaces. PMID:28163884

Aeroelastic Rotor Stability Analysis

DTIC Science & Technology

1976-01-01

1TD.l + ^2^.1 + Ac2,1Tc.i),in* + (AciV,i * ACl(-1)n<>TD.i + ^^s.l + AC3,,Tc.i,COi*:i ( B16 ) where the A’s are quantities that are Independent of...from purely dynamic or aerodynamic considerations, or both, are obtained in the same manner and can be expressed in the same form as Equation Bl6 ...If each of the rotating system equations, expressed in the form of Equation Bl6 , is multiplied by (-l)n and sunned over the number of blades N

Numerical Study of Boundary-Layer in Aerodynamics

NASA Technical Reports Server (NTRS)

Shih, Tom I-P.

1997-01-01

The accomplishments made in the following three tasks are described: (1) The first task was to study shock-wave boundary-layer interactions with bleed - this study is relevant to boundary-layer control in external and mixed-compression inlets of supersonic aircraft; (2) The second task was to test RAAKE, a code developed for computing turbulence quantities; and (3) The third task was to compute flow around the Ames ER-2 aircraft that has been retrofitted with containers over its wings and fuselage. The appendices include two reports submitted to AIAA for publication.

Fundamental aerodynamic characteristics of delta wings with leading-edge vortex flows

NASA Technical Reports Server (NTRS)

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

1985-01-01

An investigation of the aerodynamics of sharp leading-edge delta wings at supersonic speeds has been conducted. The supporting experimental data for this investigation were taken from published force, pressure, and flow-visualization data in which the Mach number normal to the wing leading edge is always less than 1.0. The individual upper- and lower-surface nonlinear characteristics for uncambered delta wings are determined and presented in three charts. The upper-surface data show that both the normal-force coefficient and minimum pressure coefficient increase nonlinearly with a decreasing slope with increasing angle of attack. The lower-surface normal-force coefficient was shown to be independent of Mach number and to increase nonlinearly, with an increasing slope, with increasing angle of attack. These charts are then used to define a wing-design space for sharp leading-edge delta wings.

An Experimental Study of the Aerodynamics of a Swept and Unswept Semispan Wing with a Simulated Glaze Ice Accretion

NASA Technical Reports Server (NTRS)

Bragg, Michael B.

1994-01-01

Two semispan wings, one with a rectangular planform and one with 30 degrees of leading edge sweep were tested. Both had a NACA 0012 airfoil section, and both were tested clean and with simulated glaze ice shapes on their leading edges. Several surface roughness were tested. Each model geometry is documented and each surface roughness is explained. Aerodynamic performance of the wing in the form of sectional lift and integrated three-dimensional lift is documented through pressure measurements obtained from rows of surface pressure taps placed at five span locations on the wing. For the rectangular wing, sectional drag near the midspan is obtained from wake total pressure profiles. The data is presented in tabular and graphical form and is also available on computer disk.

Physical and Chemical Processes in Turbulent Flames

DTIC Science & Technology

2015-06-23

positive aerodynamics stretch, into a multitude of wrinkled flamelets possessing either positive or negative stretch, such that the intensified...flame surface, such as the flame surface area ratio, build up this global measure. The turbulent flame surface is typically highly wrinkled and folded...consider a filtered/average location of the flame positions to represent a smooth surface. The information contained in the wrinkled surface if

The Effect of Negative Dihedral, Tip Droop, and Wing-tip Shape on the Low-speed Aerodynamic Characteristics of a Complete Model Having a 45 Degrees Sweptback Wing

NASA Technical Reports Server (NTRS)

Spearman, M Leroy; Becht, Robert E

1948-01-01

An investigation has been conducted in the Langley 300 MPH 7- by 10-foot tunnel to determine the effect of negative dihedral, tip droop, and wing-tip shape on the low-speed aerodynamic characteristics of a complete model having a 45 degrees sweptback wing. Longitudinal and lateral stability characteristics were obtained for the model with and without tail surfaces.

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.

A computational system for aerodynamic design and analysis of supersonic aircraft. Part 2: User's manual

NASA Technical Reports Server (NTRS)

Middleton, W. D.; Lundry, J. L.; Coleman, R. G.

1976-01-01

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. This user's manual contains a description of the system, an explanation of its usage, the input definition, and example output.

Design of a 1500 Ft/Sec, Transonic, High-through-Flow, Single-Stage Axial-Flow Compressor with Low Hub/Tip Ratio

DTIC Science & Technology

1976-10-01

aerodynamic flow field pertaining to the design point is defined on twenty-one stream surfaces, and radial and meridional distributions of significant...full radial equilibrium analysis of the compressor flow field using the streamline curvature solution technique. Through a series of iterations, it...one can assume the blade geometry, solving for the equilibriwn flow field using specified relative flow aigles as input to the aerodynamic program. In

Aeroelasticity in Turbomachines. Comparison of Theoretical and Experimental Cascade Results.

DTIC Science & Technology

1986-01-01

Y~x)csn1#(x)) It should be noted here that, in computing the blade surface pressure distribution, only components, and not amplitudes or phase angles...oscillation, done on the system is obtained by computing -i ChV+Cc+rhU+c_,Vh (12) Expressed in this way, the aerodynamic work coefficients c., cVh, cva...predictions), so the aerodynamic damping coefficient can easily be computed and plotted. This information is useful to the turbomachine designer for

Reducing the Effect of Transducer Mount Induced Noise (XMIN) on Aeroacoustic Wind Tunnel Testing Data with a New Transducer Mount Design

NASA Technical Reports Server (NTRS)

Herron, Andrew J.; Reed, Darren K.; Nance, Donald K.

2015-01-01

Characterization of flight vehicle unsteady aerodynamics is often studied via large scale wind tunnel testing. Boundary layer noise is measured by miniature pressure transducers installed in a model. Noise levels (2-5 dB ref. 20 µPa) can be induced when transducer is mounted out of flush with model outer surface. This effect must be minimized to accurately determine aerodynamically induced acoustic environments.

Hypersonic aerothermodynamic and scramjet research using high enthalpy shock tunnel

NASA Astrophysics Data System (ADS)

Itoh, K.; Ueda, S.; Tanno, H.; Komuro, T.; Sato, K.

A high enthalpy shock tunnel is a potential facility for gaining knowledge to develop modern aerothermodynamic and propulsion technologies. The largest high enthalpy shock tunnel HIEST was built at NAL Kakuda in 1997, aiming for aerothermodynamic tests of Japan's space vehicle HOPE and scramjet propulsion systems. Selected topics from the experimental studies carried out using HIEST so far, such as the nonequilibrium aerodynamics of HOPE, the surface catalytic effect on aerodynamic heating and scramjet performance are described.

Air Force Academy Aeronautics Digest - Fall/Winter 1980.

DTIC Science & Technology

1981-05-01

Crandall # _2EXAMINING A RULE OF THUMB FOR THE RELATION BETWEEN CAMBER AND 21 ZERO -LIFT ANGLE OF ATTACK,S----E.J. Jumper / EXPERIMENTAL AERODYNAMIC...slow- ing the fluid velocity to zero without loss. Static pressure is the pressure exerted on an aerodynamic surface parallel to the free stream...it is zero at the vor- tex center. Figure 2 shows the velocity distribution of a vortex with a viscous core 0. rt r Figure 2. Fluid Velocity Versus

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.

Wind-tunnel studies of the effects of simulated damage on the aerodynamic characteristics of airplanes and missiles

NASA Technical Reports Server (NTRS)

Spearman, M. L.

1979-01-01

In order to assess the effects on static aerodynamic characteristics of battle damage to an aircraft or missile, wind tunnel studies were performed on models from which all or parts of the wing or horizontal or vertical tail had been removed. The effects of damage on the lift, longitudinal stability, lateral stability and directional stability of a swept-wing fighter are presented, along with the effects of wing removal on the control requirements of a delta-wing fighter. Results indicate that the loss of a major part of the vertical tail will probably result in the loss of the aircraft at any speed, while the loss of major parts of the horizontal tail generally results in catastrophic instability at subsonic speeds but, at low supersonic speeds, may allow the aircraft to return to friendly territory before pilot ejection. Major damage to the wing may be sustained without the loss of aircraft or pilot. The loss of some of the aerodynamic surfaces of cruise or surface-to-air missiles may result in catastrophic instability or may permit a ballistic trajectory to be maintained, depending upon the location of the lost surface with respect to the center of gravity of the missile.

An experimental study of airfoil-spoiler aerodynamics

NASA Technical Reports Server (NTRS)

Mclachlan, B. G.; Karamcheti, K.

1985-01-01

The steady/unsteady flow field generated by a typical two dimensional airfoil with a statically deflected flap type spoiler was investigated. Subsonic wind tunnel tests were made over a range of parameters: spoiler deflection, angle of attack, and two Reynolds numbers; and comprehensive measurements of the mean and fluctuating surface pressures, velocities in the boundary layer, and velocities in the wake. Schlieren flow visualization of the near wake structure was performed. The mean lift, moment, and surface pressure characteristics are in agreement with previous investigations of spoiler aerodynamics. At large spoiler deflections, boundary layer character affects the static pressure distribution in the spoiler hingeline region; and, the wake mean velocity fields reveals a closed region of reversed flow aft of the spoiler. It is shown that the unsteady flow field characteristics are as follows: (1) the unsteady nature of the wake is characterized by vortex shedding; (2) the character of the vortex shedding changes with spoiler deflection; (3) the vortex shedding characteristics are in agreement with other bluff body investigations; and (4) the vortex shedding frequency component of the fluctuating surface pressure field is of appreciable magnitude at large spoiler deflections. The flow past an airfoil with deflected spoiler is a particular problem in bluff body aerodynamics is considered.

Results From F-18B Stability and Control Parameter Estimation Flight Tests at High Dynamic Pressures

NASA Technical Reports Server (NTRS)

Moes, Timothy R.; Noffz, Gregory K.; Iliff, Kenneth W.

2000-01-01

A maximum-likelihood output-error parameter estimation technique has been used to obtain stability and control derivatives for the NASA F-18B Systems Research Aircraft. This work has been performed to support flight testing of the active aeroelastic wing (AAW) F-18A project. The goal of this research is to obtain baseline F-18 stability and control derivatives that will form the foundation of the aerodynamic model for the AAW aircraft configuration. Flight data have been obtained at Mach numbers between 0.85 and 1.30 and at dynamic pressures ranging between 600 and 1500 lbf/sq ft. At each test condition, longitudinal and lateral-directional doublets have been performed using an automated onboard excitation system. The doublet maneuver consists of a series of single-surface inputs so that individual control-surface motions cannot be correlated with other control-surface motions. Flight test results have shown that several stability and control derivatives are significantly different than prescribed by the F-18B aerodynamic model. This report defines the parameter estimation technique used, presents stability and control derivative results, compares the results with predictions based on the current F-18B aerodynamic model, and shows improvements to the nonlinear simulation using updated derivatives from this research.

Catalysis study for space shuttle vehicle thermal protection systems. [for vehicle surface

NASA Technical Reports Server (NTRS)

Breen, J.; Rosner, D. E.; Delgass, W. N.; Nordine, P. C.; Cibrian, R.; Krishnan, N. G.

1973-01-01

Experimental results on the problem of reducing aerodynamic heating on space shuttle orbiter surfaces are presented. Data include: (1) development of a laboratory flow reactor technique for measuring gamma sub O and gamma sub N on candidate materials at surfaces, T sub w, in the nominal range 1000 to 2000, (2) measurements of gamma sub O and gamma sub N above 1000 K for both the glass coating of a reusable surface insulation material and the siliconized surface of a reinforced pyrolyzed plastic material, (3) measurement of the ablation behavior of the coated RPP material at T sub w is greater than or equal to 2150 K, (4) X-ray photoelectron spectral studies of the chemical constituents on these surfaces before and after dissociated gas exposure, (5) scanning electron micrograph examination of as-received and reacted specimens, and (6) development and exploitation of a method of predicting the aerodynamic heating consquences of these gamma sub O(T sub w) and gamma sub N(T sub w) measurements for critical locations on a radiation cooled orbiter vehicle.

Assessing Near-surface Heat, Water Vapor and Carbon Dioxide Exchange Over a Coastal Salt-marsh

NASA Astrophysics Data System (ADS)

Bogoev, I.; O'Halloran, T. L.; LeMoine, J.

2017-12-01

Coastal ecosystems play an important role in mitigating the effects of climate change by storing significant quantities of carbon. A growing number of studies suggest that vegetated estuarine habitats, specifically salt marshes, have high long-term rates of carbon sequestration, perhaps even higher than mature tropical and temperate forests. Large amounts of carbon, accumulated over thousands of years, are stored in the plant materials and sediment. Improved understanding of the factors that control energy and carbon exchange is needed to better guide restoration and conservation management practices. To that end, we recently established an observation system to study marsh-atmosphere interactions within the North Inlet-Winyah Bay National Estuarine Research Reserve. Near-surface fluxes of heat, water vapor (H2O) and carbon dioxide (CO2) were measured by an eddy-covariance system consisting of an aerodynamic open-path H2O / CO2 gas analyzer with a spatially integrated 3D sonic anemometer/thermometer (IRGASON). The IRGASON instrument provides co-located and highly synchronized, fast response H2O, CO2 and air- temperature measurements, which eliminates the need for spectral corrections associated with the separation between the sonic anemometer and the gas analyzer. This facilitates calculating the instantaneous CO2 molar mixing ratio relative to dry air. Fluxes computed from CO2 and H2O mixing ratios, which are conserved quantities, do not require post-processing corrections for air-density changes associated with temperature and water vapor fluctuations. These corrections are particularly important for CO2, because they could be even larger than the measured flux. Here we present the normalized frequency spectra of air temperature, water vapor and CO2, as well as their co-spectra with the co-located vertical wind. We also show mean daily cycles of sensible, latent and CO2 fluxes and analyze correlations with air/water temperature, wind speed and light availability.

On the application of a hairpin vortex model of wall turbulence to trailing edge noise prediction

NASA Technical Reports Server (NTRS)

Liu, N. S.; Shamroth, S. J.

1985-01-01

The goal is to develop a technique via a hairpin vortex model of the turbulent boundary layer, which would lead to the estimation of the aerodynamic input for use in trailing edge noise prediction theories. The work described represents an initial step in reaching this goal. The hairpin vortex is considered as the underlying structure of the wall turbulence and the turbulent boundary layer is viewed as an ensemble of typical hairpin vortices of different sizes. A synthesis technique is examined which links the mean flow and various turbulence quantities via these typical vortices. The distribution of turbulence quantities among vortices of different scales follows directly from the probability distribution needed to give the measured mean flow vorticity. The main features of individual representative hairpin vortices are discussed in detail and a preliminary assessment of the synthesis approach is made.

Experimental and Theoretical Study of Propeller Spinner/Shank Interference. M.S. Thesis

NASA Technical Reports Server (NTRS)

Cornell, C. C.

1986-01-01

A fundamental experimental and theoretical investigation into the aerodynamic interference associated with propeller spinner and shank regions was conducted. The research program involved a theoretical assessment of solutions previously proposed, followed by a systematic experimental study to supplement the existing data base. As a result, a refined computational procedure was established for prediction of interference effects in terms of interference drag and resolved into propeller thrust and torque components. These quantities were examined with attention to engineering parameters such as two spinner finess ratios, three blade shank forms, and two/three/four/six/eight blades. Consideration of the physics of the phenomena aided in the logical deduction of two individual interference quantities (cascade effects and spinner/shank juncture interference). These interference effects were semi-empirically modeled using existing theories and placed into a compatible form with an existing propeller performance scheme which provided the basis for examples of application.

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.

Formulation of aerodynamic prediction techniques for hypersonic configuration design

NASA Technical Reports Server (NTRS)

1979-01-01

An investigation of approximate theoretical techniques for predicting aerodynamic characteristics and surface pressures for relatively slender vehicles at moderate hypersonic speeds was performed. Emphasis was placed on approaches that would be responsive to preliminary configuration design level of effort. Supersonic second order potential theory was examined in detail to meet this objective. Shock layer integral techniques were considered as an alternative means of predicting gross aerodynamic characteristics. Several numerical pilot codes were developed for simple three dimensional geometries to evaluate the capability of the approximate equations of motion considered. Results from the second order computations indicated good agreement with higher order solutions and experimental results for a variety of wing like shapes and values of the hypersonic similarity parameter M delta approaching one.

Results of investigations on a 0.0405 scale model ATP version of the NR-SSV orbiter in the North American Aeronautical Laboratory low speed wind tunnel

NASA Technical Reports Server (NTRS)

Mennell, R.; Vaughn, J. E.; Singellton, R.

1973-01-01

Experimental aerodynamic investigations were conducted on a scale model space shuttle vehicle (SSV) orbiter. The purpose of the test was to investigate the longitudinal and lateral-directional aerodynamic characteristics. Emphasis was placed on model component, wing-glove, and wing-body fairing effects, as well as elevon, aileron, and rudder control effectiveness. Angles of attack from - 5 deg to + 30 deg and angles of sideslip from - 5 deg to + 10 deg were tested. Static pressures were recorded on base, fuselage, and wing surfaces. Tufts and talc-kerosene flow visualization techniques were also utilized. The aerodynamic force balance results are presented in plotted and tabular form.

Thermo-aerodynamic efficiency of non-circular ducts with vortex enhancement of heat exchange in different types of compact heat exchangers

NASA Astrophysics Data System (ADS)

Vasilev, V. Ya; Nikiforova, S. A.

2018-03-01

Experimental studies of thermo-aerodynamic characteristics of non-circular ducts with discrete turbulators on walls and interrupted channels have confirmed the rational enhancement of convective heat transfer, in which the growth of heat transfer outstrips or equals the growth of aerodynamic losses. Determining the regularities of rational (energy-saving) enhancement of heat transfer and the proposed method for comparing the characteristics of smooth-channel (without enhancement) heat exchangers with effective analogs provide new results, confirming the high efficiency of vortex enhancement of convective heat transfer in non-circular ducts of plate-finned heat exchange surfaces. This allows creating heat exchangers with much smaller mass and volume for operation in energy-saving modes.

Comprehensive helicopter rotor instrumentation - A retrofit approach using miniature transducers

NASA Technical Reports Server (NTRS)

Jacklin, Stephen A.; Mort, Ray; Morrison, Dwayne

1992-01-01

This paper reports an approach used to retrofit a set of full-scale main rotor blades with 290 miniature pressure transducers, 46 strain gages, and 24 miniature accelerometers. Normally, in order to avoid disturbing the aerodynamics of the rotor flow field, the pressure instrumentation must be integrally built into the body of the rotor blades. However, using a method developed with NASA, miniature pressure transducers are mounted to the blade exterior surface without degrading the quality of the blade aerodynamics. Moreover, it is estimated that this approach reduced costs by more than 50 percent over building a set of pressure instrumented blades. The aerodynamic measurement objectives are presented as are instrumentation design considerations, type of instrumentation used, assembly process, and the installed instrumentation characteristics.

Flight of Sharovipteryx mirabilis: the world's first delta-winged glider.

PubMed

Dyke, G J; Nudds, R L; Rayner, J M V

2006-07-01

The 225 million-year-old reptile Sharovipteryx mirabilis was the world's first delta-winged glider; this remarkable animal had a flight surface composed entirely of a hind-limb membrane. We use standard delta-wing aerodynamics to reconstruct the flight of S. mirabilis demonstrating that wing shape could have been controlled simply by protraction of the femora at the knees, and by variation in incidence of a small forelimb canard. Our method has allowed us to address the question of how identifying realistic glide performance can be used to set limits on aerodynamic design in this small animal. Our novel interpretation of the bizarre flight mode of S. mirabilis is the first based directly on interpretation of the fossil itself and the first grounded in aerodynamics.

Ground/Flight Correlation of Aerodynamic Loads with Structural Response

NASA Technical Reports Server (NTRS)

Mangalam, Arun S.; Davis, Mark C.

2009-01-01

Ground and flight tests provide a basis and methodology for in-flight characterization of the aerodynamic and structural performance through the monitoring of the fluid-structure interaction. The NF-15B flight tests of the Intelligent Flight Control System program provided a unique opportunity to test the correlation of aerodynamic loads with points of flow attaching and detaching from the surface, which are also known as flow bifurcation points, as observed in a previous wind tunnel test performed at the U.S. Air Force Academy (Colorado Springs, Colorado). Moreover, flight tests, along with the subsequent unsteady aerodynamic tests in the NASA Transonic Dynamics Tunnel (TDT), provide a basis using surface flow sensors as means of assessing the aeroelastic performance of flight vehicles. For the flight tests, the NF-15B tail was instrumented with hot-film sensors and strain gages for measuring root-bending strains. This data were gathered via selected sideslip maneuvers performed at level flight and subsonic speeds. The aerodynamic loads generated by the sideslip maneuver resulted in a structural response, which were then compared with the hot-film sensor signals. The hot-film sensor signals near the stagnation region were found to be highly correlated with the root-bending strains. For the TDT tests, a flexible wing section developed under the U.S. Air Force Research Lab SensorCraft program was instrumented with strain gages, accelerometers, and hot-film sensors at two span stations. The TDT tests confirmed the correlation between flow bifurcation points and the wing structural response to tunnel-generated gusts. Furthermore, as the wings structural modes were excited by the gusts, a gradual phase change between the flow bifurcation point and the structural mode occurred during a resonant condition.

Aeroacoustic Study of a High-Fidelity Aircraft Model: Part 1- Steady Aerodynamic Measurements

NASA Technical Reports Server (NTRS)

Khorrami, Mehdi R.; Hannon, Judith A.; Neuhart, Danny H.; Markowski, Gregory A.; VandeVen, Thomas

2012-01-01

In this paper, we present steady aerodynamic measurements for an 18% scale model of a Gulfstream air-craft. The high fidelity and highly-instrumented semi-span model was developed to perform detailed aeroacoustic studies of airframe noise associated with main landing gear/flap components and gear-flap interaction noise, as well as to evaluate novel noise reduction concepts. The aeroacoustic tests, being conducted in the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel, are split into two entries. The first entry, completed November 2010, was entirely devoted to the detailed mapping of the aerodynamic characteristics of the fabricated model. Flap deflections of 39?, 20?, and 0? with the main landing gear on and off were tested at Mach numbers of 0.16, 0.20, and 0.24. Additionally, for each flap deflection, the model was tested with the tunnel both in the closed-wall and open-wall (jet) modes. During this first entry, global forces (lift and drag) and extensive steady and unsteady surface pressure measurements were obtained. Preliminary analysis of the measured forces indicates that lift, drag, and stall characteristics compare favorably with Gulfstream?s high Reynolds number flight data. The favorable comparison between wind-tunnel and flight data allows the semi-span model to be used as a test bed for developing/evaluating airframe noise reduction concepts under a relevant environment. Moreover, initial comparison of the aerodynamic measurements obtained with the tunnel in the closed- and open-wall configurations shows similar aerodynamic behavior. This permits the acoustic and off-surface flow measurements, planned for the second entry, to be conducted with the tunnel in the open-jet mode.

The effects of inlet turbulence and rotor/stator interactions on the aerodynamics and heat transfer of a large-scale rotating turbine model. Part 4: Aerodynamic data tabulation

NASA Technical Reports Server (NTRS)

Dring, R. P.; Joslyn, H. D.; Blair, M. F.

1987-01-01

A combined experimental and analytical program was conducted to examine the effects of inlet turbulence and airfoil heat transfer. The experimental portion of the study was conducted in a large-scale (approx. 5X engine), ambient temperature, rotating turbine model configured in both single-stage and stage-and-a-half arrangements. Heat transfer measurements were obtained using low-conductivity airfoils with miniature thermocouples welded to a thin, electrically heated surface skin. Heat transfer data were acquired for various combinations of low or high inlet turbulence intensity, flow coefficient, first stator-rotor axial spacing, Reynolds number and relative circumferential position of the first and second stators. Aerodynamic measurements obtained include distributions of the mean and fluctuating velocities at the turbine inlet and, for each airfoil row, midspan airfoil surface pressures and circumferential distributions of the downstream steady state pressures and fluctuating velocities. Results include airfoil heat transfer predictions produced using existing 2-D boundary layer computation schemes and an examination of solutions of the unsteady boundary layer equations.

Unstructured Grid Euler Method Assessment for Longitudinal and Lateral/Directional Aerodynamic Performance Analysis of the HSR Technology Concept Airplane at Supersonic Cruise Speed

NASA Technical Reports Server (NTRS)

Ghaffari, Farhad

1999-01-01

Unstructured grid Euler computations, performed at supersonic cruise speed, are presented for a High Speed Civil Transport (HSCT) configuration, designated as the Technology Concept Airplane (TCA) within the High Speed Research (HSR) Program. The numerical results are obtained for the complete TCA cruise configuration which includes the wing, fuselage, empennage, diverters, and flow through nacelles at M (sub infinity) = 2.4 for a range of angles-of-attack and sideslip. Although all the present computations are performed for the complete TCA configuration, appropriate assumptions derived from the fundamental supersonic aerodynamic principles have been made to extract aerodynamic predictions to complement the experimental data obtained from a 1.675%-scaled truncated (aft fuselage/empennage components removed) TCA model. The validity of the computational results, derived from the latter assumptions, are thoroughly addressed and discussed in detail. The computed surface and off-surface flow characteristics are analyzed and the pressure coefficient contours on the wing lower surface are shown to correlate reasonably well with the available pressure sensitive paint results, particularly, for the complex flow structures around the nacelles. The predicted longitudinal and lateral/directional performance characteristics for the truncated TCA configuration are shown to correlate very well with the corresponding wind-tunnel data across the examined range of angles-of-attack and sideslip. The complementary computational results for the longitudinal and lateral/directional performance characteristics for the complete TCA configuration are also presented along with the aerodynamic effects due to empennage components. Results are also presented to assess the computational method performance, solution sensitivity to grid refinement, and solution convergence characteristics.

Use of the Ames Check Standard Model for the Validation of Wall Interference Corrections

NASA Technical Reports Server (NTRS)

Ulbrich, N.; Amaya, M.; Flach, R.

2018-01-01

The new check standard model of the NASA Ames 11-ft Transonic Wind Tunnel was chosen for a future validation of the facility's wall interference correction system. The chosen validation approach takes advantage of the fact that test conditions experienced by a large model in the slotted part of the tunnel's test section will change significantly if a subset of the slots is temporarily sealed. Therefore, the model's aerodynamic coefficients have to be recorded, corrected, and compared for two different test section configurations in order to perform the validation. Test section configurations with highly accurate Mach number and dynamic pressure calibrations were selected for the validation. First, the model is tested with all test section slots in open configuration while keeping the model's center of rotation on the tunnel centerline. In the next step, slots on the test section floor are sealed and the model is moved to a new center of rotation that is 33 inches below the tunnel centerline. Then, the original angle of attack sweeps are repeated. Afterwards, wall interference corrections are applied to both test data sets and response surface models of the resulting aerodynamic coefficients in interference-free flow are generated. Finally, the response surface models are used to predict the aerodynamic coefficients for a family of angles of attack while keeping dynamic pressure, Mach number, and Reynolds number constant. The validation is considered successful if the corrected aerodynamic coefficients obtained from the related response surface model pair show good agreement. Residual differences between the corrected coefficient sets will be analyzed as well because they are an indicator of the overall accuracy of the facility's wall interference correction process.

An Aeroacoustic Characterization of a Multi-Element High-Lift Airfoil

NASA Astrophysics Data System (ADS)

Pascioni, Kyle A.

The leading edge slat of a high-lift system is known to be a large contributor to the overall radiated acoustic field from an aircraft during the approach phase of the flight path. This is due to the unsteady flow field generated in the slat-cove and near the leading edge of the main element. In an effort to understand the characteristics of the flow-induced source mechanisms, a suite of experimental measurements has been performed on a two-dimensional multi-element airfoil, namely, the MD-30P30N. Particle image velocimetry provide mean flow field and turbulence statistics to illustrate the differences associated with a change in angle of attack. Phase-averaged quantities prove shear layer instabilities to be linked to narrowband peaks found in the acoustic spectrum. Unsteady surface pressure are also acquired, displaying strong narrowband peaks and large spanwise coherence at low angles of attack, whereas the spectrum becomes predominately broadband at high angles. Nonlinear frequency interaction is found to occur at low angles of attack, while being negligible at high angles. To localize and quantify the noise sources, phased microphone array measurements are per- formed on the two dimensional high-lift configuration. A Kevlar wall test section is utilized to allow the mean aerodynamic flow field to approach distributions similar to a free-air configuration, while still capable of measuring the far field acoustic signature. However, the inclusion of elastic porous sidewalls alters both aerodynamic and acoustic characteristics. Such effects are considered and accounted for. Integrated spectra from Delay and Sum and DAMAS beamforming effectively suppress background facility noise and additional noise generated at the tunnel wall/airfoil junction. Finally, temporally-resolved estimates of a low-dimensional representation of the velocity vector fields are obtained through the use of proper orthogonal decomposition and spectral linear stochastic estimation. An estimate of the pressure field is then extracted by Poissons equation. From this, Curles analogy projects the time-resolved pressure forces on the airfoil surface to further establish the connection between the dominating unsteady flow structures and the propagated noise.

Scramjet exhaust simulation technique for hypersonic aircraft nozzle design and aerodynamic tests

NASA Technical Reports Server (NTRS)

Hunt, J. L.; Talcott, N. A., Jr.; Cubbage, J. M.

1977-01-01

Current design philosophy for scramjet-powered hypersonic aircraft results in configurations with the entire lower fuselage surface utilized as part of the propulsion system. The lower aft-end of the vehicle acts as a high expansion ratio nozzle. Not only must the external nozzle be designed to extract the maximum possible thrust force from the high energy flow at the combustor exit, but the forces produced by the nozzle must be aligned such that they do not unduly affect aerodynamic balance. The strong coupling between the propulsion system and aerodynamics of the aircraft makes imperative at least a partial simulation of the inlet, exhaust, and external flows of the hydrogen-burning scramjet in conventional facilities for both nozzle formulation and aerodynamic-force data acquisition. Aerodynamic testing methods offer no contemporary approach for such vehicle design requirements. NASA-Langley has pursued an extensive scramjet/airframe integration R&D program for several years and has recently developed a promising technique for simulation of the scramjet exhaust flow for hypersonic aircraft. Current results of the research program to develop a scramjet flow simulation technique through the use of substitute gas blends are described in this paper.

Aerodynamics of Race Cars

NASA Astrophysics Data System (ADS)

Katz, Joseph

2006-01-01

Race car performance depends on elements such as the engine, tires, suspension, road, aerodynamics, and of course the driver. In recent years, however, vehicle aerodynamics gained increased attention, mainly due to the utilization of the negative lift (downforce) principle, yielding several important performance improvements. This review briefly explains the significance of the aerodynamic downforce and how it improves race car performance. After this short introduction various methods to generate downforce such as inverted wings, diffusers, and vortex generators are discussed. Due to the complex geometry of these vehicles, the aerodynamic interaction between the various body components is significant, resulting in vortex flows and lifting surface shapes unlike traditional airplane wings. Typical design tools such as wind tunnel testing, computational fluid dynamics, and track testing, and their relevance to race car development, are discussed as well. In spite of the tremendous progress of these design tools (due to better instrumentation, communication, and computational power), the fluid dynamic phenomenon is still highly nonlinear, and predicting the effect of a particular modification is not always trouble free. Several examples covering a wide range of vehicle shapes (e.g., from stock cars to open-wheel race cars) are presented to demonstrate this nonlinear nature of the flow field.

Aerothermal Analysis and Design of the Gravity Recovery and Climate Experiment (GRACE) Spacecraft

NASA Technical Reports Server (NTRS)

Mazanek, Daniel D.; Kumar, Renjith R.; Qu, Min; Seywald, Hans

2000-01-01

The Gravity Recovery and Climate Experiment (GRACE) primary mission will be performed by making measurements of the inter-satellite range change between two co-planar, low altitude near-polar orbiting satellites. Understanding the uncertainties in the disturbance environment, particularly the aerodynamic drag and torques, is critical in several mission areas. These include an accurate estimate of the spacecraft orbital lifetime, evaluation of spacecraft attitude control requirements, and estimation of the orbital maintenance maneuver frequency necessitated by differences in the drag forces acting on both satellites. The FREEMOL simulation software has been developed and utilized to analyze and suggest design modifications to the GRACE spacecraft. Aerodynamic accommodation bounding analyses were performed and worst-case envelopes were obtained for the aerodynamic torques and the differential ballistic coefficients between the leading and trailing GRACE spacecraft. These analyses demonstrate how spacecraft aerodynamic design and analysis can benefit from a better understanding of spacecraft surface accommodation properties, and the implications for mission design constraints such as formation spacing control.

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.

Simulated flight acoustic investigation of treated ejector effectiveness on advanced mechanical suppresors for high velocity jet noise reduction

NASA Technical Reports Server (NTRS)

Brausch, J. F.; Motsinger, R. E.; Hoerst, D. J.

1986-01-01

Ten scale-model nozzles were tested in an anechoic free-jet facility to evaluate the acoustic characteristics of a mechanically suppressed inverted-velocity-profile coannular nozzle with an accoustically treated ejector system. The nozzle system used was developed from aerodynamic flow lines evolved in a previous contract, defined to incorporate the restraints imposed by the aerodynamic performance requirements of an Advanced Supersonic Technology/Variable Cycle Engine system through all its mission phases. Accoustic data of 188 test points were obtained, 87 under static and 101 under simulated flight conditions. The tests investigated variables of hardwall ejector application to a coannular nozzle with 20-chute outer annular suppressor, ejector axial positioning, treatment application to ejector and plug surfaces, and treatment design. Laser velocimeter, shadowgraph photograph, aerodynamic static pressure, and temperature measurement were acquired on select models to yield diagnositc information regarding the flow field and aerodynamic performance characteristics of the nozzles.

GRACE Mission Design: Impact of Uncertainties in Disturbance Environment and Satellite Force Models

NASA Technical Reports Server (NTRS)

Mazanek, Daniel D.; Kumar, Renjith R.; Seywald, Hans; Qu, Min

2000-01-01

The Gravity Recovery and Climate Experiment (GRACE) primary mission will be performed by making measurements of the inter-satellite range change between two co-planar, low altitude, near-polar orbiting satellites. Understanding the uncertainties in the disturbance environment, particularly the aerodynamic drag and torques, is critical in several mission areas. These include an accurate estimate of the spacecraft orbital lifetime, evaluation of spacecraft attitude control requirements, and estimation of the orbital maintenance maneuver frequency necessitated by differences in the drag forces acting on both satellites. The FREEMOL simulation software has been developed and utilized to analyze and suggest design modifications to the GRACE spacecraft. Aerodynamic accommodation bounding analyses were performed and worst-case envelopes were obtained for the aerodynamic torques and the differential ballistic coefficients between the leading and trailing GRACE spacecraft. These analyses demonstrate how spacecraft aerodynamic design and analysis can benefit from a better understanding of spacecraft surface accommodation properties, and the implications for mission design constraints such as formation spacing control.

A new specimen of the Early Cretaceous bird Hongshanornis longicresta: insights into the aerodynamics and diet of a basal ornithuromorph.

PubMed

Chiappe, Luis M; Zhao, Bo; O'Connor, Jingmai K; Chunling, Gao; Wang, Xuri; Habib, Michael; Marugan-Lobon, Jesus; Meng, Qingjin; Cheng, Xiaodong

2014-01-01

The discovery of Hongshanornis longicresta, a small ornithuromorph bird with unusually long hindlimb proportions, was followed by the discovery of two closely related species, Longicrusavis houi and Parahongshanornis chaoyangensis. Together forming the Hongshanornithidae, these species reveal important information about the early diversity and morphological specialization of ornithuromorphs, the clade that contains all living birds. Here we report on a new specimen (DNHM D2945/6) referable to Hongshanornis longicresta that contributes significant information to better understand the morphology, trophic ecology, and aerodynamics of this species, as well as the taxonomy of the Hongshanornithidae. Most notable are the well-preserved wings and feathered tail of DNHM D2945/6, which afford an accurate reconstruction of aerodynamic parameters indicating that as early as 125 million years ago, basal ornithuromorphs had evolved aerodynamic surfaces comparable in size and design to those of many modern birds, and flight modes alike to those of some small living birds.

A new specimen of the Early Cretaceous bird Hongshanornis longicresta: insights into the aerodynamics and diet of a basal ornithuromorph

PubMed Central

Zhao, Bo; O’Connor, Jingmai K.; Chunling, Gao; Wang, Xuri; Habib, Michael; Marugan-Lobon, Jesus; Meng, Qingjin; Cheng, Xiaodong

2014-01-01

The discovery of Hongshanornis longicresta, a small ornithuromorph bird with unusually long hindlimb proportions, was followed by the discovery of two closely related species, Longicrusavis houi and Parahongshanornis chaoyangensis. Together forming the Hongshanornithidae, these species reveal important information about the early diversity and morphological specialization of ornithuromorphs, the clade that contains all living birds. Here we report on a new specimen (DNHM D2945/6) referable to Hongshanornis longicresta that contributes significant information to better understand the morphology, trophic ecology, and aerodynamics of this species, as well as the taxonomy of the Hongshanornithidae. Most notable are the well-preserved wings and feathered tail of DNHM D2945/6, which afford an accurate reconstruction of aerodynamic parameters indicating that as early as 125 million years ago, basal ornithuromorphs had evolved aerodynamic surfaces comparable in size and design to those of many modern birds, and flight modes alike to those of some small living birds. PMID:24482756

Longitudinal aerodynamic characteristics of a wing-winglet model designed at M = 0.8, C sub L = 0.4 using linear aerodynamic theory

NASA Technical Reports Server (NTRS)

Kuhlman, J. M.

1983-01-01

Wind tunnel test results have been presented herein for a subsonic transport type wing fitted with winglets. Wind planform was chosen to be representative of wings used on current jet transport aircraft, while wing and winglet camber surfaces were designed using two different linear aerodynamic design methods. The purpose of the wind tunnel investigation was to determine the effectiveness of these linear aerodynamic design computer codes in designing a non-planar transport configuration which would cruise efficiently. The design lift coefficient was chosen to be 0.4, at a design Mach number of 0.8. Force and limited pressure data were obtained for the basic wing, and for the wing fitted with the two different winglet designs, at Mach numbers of 0.60, 0.70, 0.75 and 0.80 over an angle of attack range of -2 to +6 degrees, at zero sideslip. The data have been presented without analysis to expedite publication.

Tomographic particle image velocimetry of desert locust wakes: instantaneous volumes combine to reveal hidden vortex elements and rapid wake deformation

PubMed Central

Bomphrey, Richard J.; Henningsson, Per; Michaelis, Dirk; Hollis, David

2012-01-01

Aerodynamic structures generated by animals in flight are unstable and complex. Recent progress in quantitative flow visualization has advanced our understanding of animal aerodynamics, but measurements have hitherto been limited to flow velocities at a plane through the wake. We applied an emergent, high-speed, volumetric fluid imaging technique (tomographic particle image velocimetry) to examine segments of the wake of desert locusts, capturing fully three-dimensional instantaneous flow fields. We used those flow fields to characterize the aerodynamic footprint in unprecedented detail and revealed previously unseen wake elements that would have gone undetected by two-dimensional or stereo-imaging technology. Vortex iso-surface topographies show the spatio-temporal signature of aerodynamic force generation manifest in the wake of locusts, and expose the extent to which animal wakes can deform, potentially leading to unreliable calculations of lift and thrust when using conventional diagnostic methods. We discuss implications for experimental design and analysis as volumetric flow imaging becomes more widespread. PMID:22977102

The Total In-Flight Simulator (TIFS) aerodynamics and systems: Description and analysis. [maneuver control and gust alleviators

NASA Technical Reports Server (NTRS)

Andrisani, D., II; Daughaday, H.; Dittenhauser, J.; Rynaski, E.

1978-01-01

The aerodynamics, control system, instrumentation complement and recording system of the USAF Total In/Flight Simulator (TIFS) airplane are described. A control system that would allow the ailerons to be operated collectively, as well as, differentially to entrance the ability of the vehicle to perform the dual function of maneuver load control and gust alleviation is emphasized. Mathematical prediction of the rigid body and the flexible equations of longitudinal motion using the level 2.01 FLEXSTAB program are included along with a definition of the vehicle geometry, the mass and stiffness distribution, the calculated mode frequencies and mode shapes, and the resulting aerodynamic equations of motion of the flexible vehicle. A complete description of the control and instrumentation system of the aircraft is presented, including analysis, ground test and flight data comparisons of the performance and bandwidth of the aerodynamic surface servos. Proposed modification for improved performance of the servos are also presented.

Lewis icing research tunnel test of the aerodynamic effects of aircraft ground deicing/anti-icing fluids

NASA Technical Reports Server (NTRS)

Runyan, L. James; Zierten, Thomas A.; Hill, Eugene G.; Addy, Harold E., Jr.

1992-01-01

A wind tunnel investigation of the effect of aircraft ground deicing/anti-icing fluids on the aerodynamic characteristics of a Boeing 737-200ADV airplane was conducted. The test was carried out in the NASA Lewis Icing Research Tunnel. Fluids tested include a Newtonian deicing fluid, three non-Newtonian anti-icing fluids commercially available during or before 1988, and eight new experimental non-Newtonian fluids developed by four fluid manufacturers. The results show that fluids remain on the wind after liftoff and cause a measurable lift loss and drag increase. These effects are dependent on the high-lift configuration and on the temperature. For a configuration with a high-lift leading-edge device, the fluid effect is largest at the maximum lift condition. The fluid aerodynamic effects are related to the magnitude of the fluid surface roughness, particularly in the first 30 percent chord. The experimental fluids show a significant reduction in aerodynamic effects.

Application of pneumatic lift and control surface technology to advanced transport aircraft

NASA Technical Reports Server (NTRS)

Englar, Robert J.

1996-01-01

The application of pneumatic (blown) aerodynamic technology to both the lifting and the control surfaces of advanced transport aircraft can provide revolutionary changes in the performance and operation of these vehicles, ranging in speed regime from Advanced Subsonic Transports to the High Speed Civil Transport, and beyond. This technology, much of it based on the Circulation Control Wing blown concepts, can provide aerodynamic force augmentations of 80 to 100 (i.e., return of 80-100 pounds of force per pound of input momentum from the blowing jet). This can be achieved without use of external mechanical surfaces. Clever application of this technology can provide no-moving-part lifting surfaces (wings/tails) integrated into the control system to greatly simplify aircraft designs while improving their aerodynamic performance. Lift/drag ratio may be pneumatically tailored to fit the current phase of the flight, and takeoff/landing performance can be greatly improved by reducing ground roll distances and liftoff/touchdown speeds. Alternatively, great increases in liftoff weights and payloads are possible, as are great reductions in wing and tail planform size, resulting in optimized cruise wing designs. Furthermore, lift generation independent of angle of attack provides much promise for increased safety of flight in the severe updrafts/downdrafts of microbursts and windshears, which is further augmented by the ability to sustain flight at greatly reduced airspeeds. Load-tailored blown wings can also reduce tip vorticity during highlift operations and the resulting vortex wake hazards near terminal areas. Reduced noise may also be possible as these jets can be made to operate at low pressures. The planned presentation will support the above statements through discussions of recent experimental and numerical (CFD) research and development of these advanced blown aerodynamic surfaces, portions of which have been conducted for NASA. Also to be presented will be predicted performance of advanced transports resulting from these devices. Suggestions will be presented for additional innovative high-payoff research leading to further confirmation of these concepts and their application to advanced efficient commercial transport aircraft.

On the calculation of dynamic and heat loads on a three-dimensional body in a hypersonic flow

NASA Astrophysics Data System (ADS)

Bocharov, A. N.; Bityurin, V. A.; Evstigneev, N. M.; Fortov, V. E.; Golovin, N. N.; Petrovskiy, V. P.; Ryabkov, O. I.; Teplyakov, I. O.; Shustov, A. A.; Solomonov, Yu S.

2018-01-01

We consider a three-dimensional body in a hypersonic flow at zero angle of attack. Our aim is to estimate heat and aerodynamic loads on specific body elements. We are considering a previously developed code to solve coupled heat- and mass-transfer problem. The change of the surface shape is taken into account by formation of the iterative process for the wall material ablation. The solution is conducted on the multi-graphics-processing-unit (multi-GPU) cluster. Five Mach number points are considered, namely for M = 20-28. For each point we estimate body shape after surface ablation, heat loads on the surface and aerodynamic loads on the whole body and its elements. The latter is done using Gauss-type quadrature on the surface of the body. The comparison of the results for different Mach numbers is performed. We also estimate the efficiency of the Navier-Stokes code on multi-GPU and central processing unit architecture for the coupled heat and mass transfer problem.

Effect of wing-tip dihedral on the longitudinal and lateral aerodynamic characteristics of a supersonic cruise configuration at subsonic speeds

NASA Technical Reports Server (NTRS)

Washburn, K. E.; Gloss, B. B.

1976-01-01

Force and moment data studies were conducted to determine the effect of wing-tip dihedral on the longitudinal and lateral aerodynamic characteristics of a supersonic cruise fighter configuration. Oil flow studies were also performed to investigate the model surface flow. Three models were tested: a flat (0 deg dihedral) wing tip, a dihedral, and an anhedral wing tip. The tests were conducted at the NASA Langley high-speed 7- by 10-foot wind tunnel.

A Digital Program for Calculating the Interaction Between Flexible Structures, Unsteady Aerodynamics and Active Controls

NASA Technical Reports Server (NTRS)

Peele, E. L.; Adams, W. M., Jr.

1979-01-01

A computer program, ISAC, is described which calculates the stability and response of a flexible airplane equipped with active controls. The equations of motion relative to a fixed inertial coordinate system are formulated in terms of the airplane's rigid body motion and its unrestrained normal vibration modes. Unsteady aerodynamic forces are derived from a doublet lattice lifting surface theory. The theoretical basis for the program is briefly explained together with a description of input data and output results.

Fitting aerodynamics and propulsion into the puzzle

NASA Technical Reports Server (NTRS)

Johnston, Patrick J.; Whitehead, Allen H., Jr.; Chapman, Gary T.

1987-01-01

The development of an airbreathing single-stage-to-orbit vehicle, in particular the problems of aerodynamics and propulsion integration, is examined. The boundary layer transition on constant pressure surfaces at hypersonic velocities, and the effects of noise on the transition are investigated. The importance of viscosity, real-gas effects, and drag at hypersonic speeds is discussed. A propulsion system with sufficient propulsive lift to enhance the performance of the vehicle is being developed. The difficulties of engine-airframe integration are analyzed.

State-of-the-Art Engineering Aeroprediction Methods with Emphasis on New Semiempirical Techniques for Predicting Nonlinear Aerodynamics on Complete Missile Configurations

DTIC Science & Technology

1993-11-01

ON COMPLETE MISSILE CONFIGURATIONS BY FRANK G. MOORE WEAPONS SYSTEMS DEPARTMENT NOVEMBER 1993 DTIC ,A -LECTE. DEC 1 3 1993 7;] Approved for public...Di t ib , tio Av J,fia ii !y Cc,,jes Approved for public release; distribution is unlimited. Avii j iUr Di"t NAVAL SURFACE WARFARE CENTER DAHLGREN...CONTENTS Setion 1.0 INTRODUCTION AND BACKGROUND .................... 1 - 1 1.1 USES FOR AERODYNAMICS ........................... 1 - 1 1.2 TYPES OF

The Effects of Warhead-Induced Damage on the Aeroelastic Characteristics of Lifting Surfaces. Volume 2. Aerodynamic Effects

DTIC Science & Technology

1980-07-01

port Scannivalves, each with a DRUCK PDCR22, I psid pressure transducer. MODEL The models used in all the aerodynamic tests were left-half, T-38...14 - 2. . -0. 3 c3 -G.076 -- . 7 -0. -1.047 -0.031 .01A .G,2 15 -0.0.2 -0. Wl -0.0F2 -0.’q.1 -1 .c; 0. ... I ncilence: -7. -ei. fl’ 3D ’:i~t’): 0.010 e

Wind tunnel investigation of the aerodynamic characteristics of symmetrically deflected ailerons of the F-8C airplane. [conducted in the Langley 8-foot transonic pressure tunnel

NASA Technical Reports Server (NTRS)

Gera, J.

1977-01-01

A .042-scale model of the F-8C airplane was investigated in a transonic wind tunnel at high subsonic Mach numbers and a range of angles of attack between-3 and 20 degrees. The effect of symmetrically deflected ailerons on the longitudinal aerodynamic characteristics was measured. Some data were also obtained on the lateral control effectiveness of asymmetrically deflected horizontal tail surfaces.

Numerical investigation on properties of attack angle for an opposing jet thermal protection system

NASA Astrophysics Data System (ADS)

Lu, Hai-Bo; Liu, Wei-Qiang

2012-08-01

The three-dimensional Navier—Stokes equation and the k-in viscous model are used to simulate the attack angle characteristics of a hemisphere nose-tip with an opposing jet thermal protection system in supersonic flow conditions. The numerical method is validated by the relevant experiment. The flow field parameters, aerodynamic forces, and surface heat flux distributions for attack angles of 0°, 2°, 5°, 7°, and 10° are obtained. The detailed numerical results show that the cruise attack angle has a great influence on the flow field parameters, aerodynamic force, and surface heat flux distribution of the supersonic vehicle nose-tip with an opposing jet thermal protection system. When the attack angle reaches 10°, the heat flux on the windward generatrix is close to the maximal heat flux on the wall surface of the nose-tip without thermal protection system, thus the thermal protection has failed.

Steady and Oscillatory, Subsonic and Supersonic, Aerodynamic Pressure and Generalized Forces for Complex Aircraft Configurations and Applications to Flutter. M.S. Thesis

NASA Technical Reports Server (NTRS)

Chen, L. T.

1975-01-01

A general method for analyzing aerodynamic flows around complex configurations is presented. By applying the Green function method, a linear integral equation relating the unknown, small perturbation potential on the surface of the body, to the known downwash is obtained. The surfaces of the aircraft, wake and diaphragm (if necessary) are divided into small quadrilateral elements which are approximated with hyperboloidal surfaces. The potential and its normal derivative are assumed to be constant within each element. This yields a set of linear algebraic equations and the coefficients are evaluated analytically. By using Gaussian elimination method, equations are solved for the potentials at the centroids of elements. The pressure coefficient is evaluated by the finite different method; the lift and moment coefficients are evaluated by numerical integration. Numerical results are presented, and applications to flutter are also included.

PROGRAM VSAERO: A computer program for calculating the non-linear aerodynamic characteristics of arbitrary configurations: User's manual

NASA Technical Reports Server (NTRS)

Maskew, B.

1982-01-01

VSAERO is a computer program used to predict the nonlinear aerodynamic characteristics of arbitrary three-dimensional configurations in subsonic flow. Nonlinear effects of vortex separation and vortex surface interaction are treated in an iterative wake-shape calculation procedure, while the effects of viscosity are treated in an iterative loop coupling potential-flow and integral boundary-layer calculations. The program employs a surface singularity panel method using quadrilateral panels on which doublet and source singularities are distributed in a piecewise constant form. This user's manual provides a brief overview of the mathematical model, instructions for configuration modeling and a description of the input and output data. A listing of a sample case is included.

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.

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.

Diminished tektite ablation in the wake of a swarm

NASA Technical Reports Server (NTRS)

Sepri, P.; Chen, K. K.; Okeefe, J. A.

1981-01-01

Observations of ablation markings on tektite surfaces reveal that a large variation in aerodynamic heating must have occurred among the members of a swarm during atmospheric entry. In a few cases, the existence of jagged features indicates that these tektite surfaces may have barely reached the melting temperature. Such an observation seems to be incompatible with the necessarily large heating rates suffered by other tektites which exhibit the ring wave melt flow. A reconciliation is proposed in the form of a wake shielding model which is a natural consequence of swarm entry. Calculations indicate that the observed ablation variations are actually possible for swarm entry at greater than escape velocity. This aerodynamic conclusion provides support for the arguments favoring extraterrestrial origin of tektites.

Finite state modeling of aeroelastic systems

NASA Technical Reports Server (NTRS)

Vepa, R.

1977-01-01

A general theory of finite state modeling of aerodynamic loads on thin airfoils and lifting surfaces performing completely arbitrary, small, time-dependent motions in an airstream is developed and presented. The nature of the behavior of the unsteady airloads in the frequency domain is explained, using as raw materials any of the unsteady linearized theories that have been mechanized for simple harmonic oscillations. Each desired aerodynamic transfer function is approximated by means of an appropriate Pade approximant, that is, a rational function of finite degree polynomials in the Laplace transform variable. The modeling technique is applied to several two dimensional and three dimensional airfoils. Circular, elliptic, rectangular and tapered planforms are considered as examples. Identical functions are also obtained for control surfaces for two and three dimensional airfoils.

Effect of coolant flow ejection on aerodynamic performance of low-aspect-ratio vanes. 1: Performance with coolant ejection holes plugged

NASA Technical Reports Server (NTRS)

Haas, J. E.; Kofskey, M. G.

1976-01-01

The aerodynamic performance of a low aspect ratio turbine vane designed with coolant flow ejection holes on the vane surfaces was experimentally determined in a full-annular cascade with the coolant ejection holes plugged. The purpose was to establish a baseline for comparison with tests where flow is ejected from the vane surfaces. The vanes were tested over a mean-section ideal critical velocity ratio range of 0.64 to 0.98. This ideal critical velocity ratio corresponds to the vane inlet total to vane aftermixed static pressure ratio at the mean section. The variations in vane efficiency and aftermixed flow conditions with circumferential and radial position were obtained.

Aerodynamic flight control to increase payload capability of future launch vehicles

NASA Technical Reports Server (NTRS)

Cochran, John E., Jr.; Cheng, Y.-M.; Leleux, Todd; Bigelow, Scott; Hasbrook, William

1993-01-01

In this report, we provide some examples of French, Russian, Chinese, and Japanese launch vehicles that have utilized fins in their designs. Next, the aerodynamic design of the fins is considered in Section 3. Some comments on basic static stability and control theory are followed by a brief description of an aerodynamic characteristics prediction code that was used to estimate the characteristics of a modified NLS 1.5 Stage vehicle. Alternative fin designs are proposed and some estimated aerodynamic characteristics presented and discussed. Also included in Section 3 is a discussion of possible methods of enhancement of the aerodynamic efficiency of fins, such as vortex generators and jet flaps. We consider the construction of fins for launch vehicles in Section 4 and offer an assessment of the state-of-the-art in the use of composites for aerodynamic control surfaces on high speed vehicles. We also comment on the use of smart materials for launch vehicle fins. The dynamic stability and control of a launch vehicle that utilizes both thrust vector control (engine nozzle gimballing) and movable fins is the subject addressed in Section 5. We give a short derivation of equations of motion for a launch vehicle moving in a vertical plane above a spherical earth, discuss the use of a gravity-turn nominal trajectory, and give the form of the period equations linearized about such a nominal. We then consider feedback control of vehicle attitude using both engine gimballing and fin deflection. Conclusions are stated and recommendations made in Section 6. An appendix contains aerodynamic data in tabular and graphical formats.

Aerodynamic Surface Stress Intermittency and Conditionally Averaged Turbulence Statistics

NASA Astrophysics Data System (ADS)

Anderson, W.

2015-12-01

Aeolian erosion of dry, flat, semi-arid landscapes is induced (and sustained) by kinetic energy fluxes in the aloft atmospheric surface layer. During saltation -- the mechanism responsible for surface fluxes of dust and sediment -- briefly suspended sediment grains undergo a ballistic trajectory before impacting and `splashing' smaller-diameter (dust) particles vertically. Conceptual models typically indicate that sediment flux, q (via saltation or drift), scales with imposed aerodynamic (basal) stress raised to some exponent, n, where n > 1. Since basal stress (in fully rough, inertia-dominated flows) scales with the incoming velocity squared, u^2, it follows that q ~ u^2n (where u is some relevant component of the above flow field, u(x,t)). Thus, even small (turbulent) deviations of u from its time-averaged value may play an enormously important role in aeolian activity on flat, dry landscapes. The importance of this argument is further augmented given that turbulence in the atmospheric surface layer exhibits maximum Reynolds stresses in the fluid immediately above the landscape. In order to illustrate the importance of surface stress intermittency, we have used conditional averaging predicated on aerodynamic surface stress during large-eddy simulation of atmospheric boundary layer flow over a flat landscape with momentum roughness length appropriate for the Llano Estacado in west Texas (a flat agricultural region that is notorious for dust transport). By using data from a field campaign to measure diurnal variability of aeolian activity and prevailing winds on the Llano Estacado, we have retrieved the threshold friction velocity (which can be used to compute threshold surface stress under the geostrophic balance with the Monin-Obukhov similarity theory). This averaging procedure provides an ensemble-mean visualization of flow structures responsible for erosion `events'. Preliminary evidence indicates that surface stress peaks are associated with the passage of inclined, high-momentum regions flanked by adjacent low-momentum regions. We will characterize geometric attributes of such structures and explore streamwise and vertical vorticity distribution within the conditionally averaged flow field.

Orion Flight Test-1 Thermal Protection System Instrumentation

NASA Technical Reports Server (NTRS)

Kowal, T. John

2011-01-01

The Orion Crew Exploration Vehicle (CEV) was originally under development to provide crew transport to the International Space Station after the retirement of the Space Shuttle, and to provide a means for the eventual return of astronauts to the Moon. With the current changes in the future direction of the United States human exploration programs, the focus of the Orion project has shifted to the project s first orbital flight test, designated Orion Flight Test 1 (OFT-1). The OFT-1 is currently planned for launch in July 2013 and will demonstrate the Orion vehicle s capability for performing missions in low Earth orbit (LEO), as well as extensibility beyond LEO for select, critical areas. Among the key flight test objectives are those related to validation of the re-entry aerodynamic and aerothermal environments, and the performance of the thermal protection system (TPS) when exposed to these environments. A specific flight test trajectory has been selected to provide a high energy entry beyond that which would be experienced during a typical low Earth orbit return, given the constraints imposed by the possible launch vehicles. This trajectory resulted from a trade study that considered the relative benefit of conflicting objectives from multiple subsystems, and sought to provide the maximum integrated benefit to the re-entry state-of-the-art. In particular, the trajectory was designed to provide: a significant, measureable radiative heat flux to the windward surface; data on boundary transition from laminar to turbulent flow; and data on catalytic heating overshoot on non-ablating TPS. In order to obtain the necessary flight test data during OFT-1, the vehicle will need to have an adequate quantity of instrumentation. A collection of instrumentation is being developed for integration in the OFT-1 TPS. In part, this instrumentation builds upon the work performed for the Mars Science Laboratory Entry, Descent and Landing Instrument (MEDLI) suite to instrument the OFT-1 ablative heat shield. The MEDLI integrated sensor plugs and pressure sensors will be adapted for compatibility with the Orion TPS design. The sensor plugs will provide in-depth temperature data to support aerothermal and TPS model correlation, and the pressure sensors will provide a flush air data system for validation of the entry and descent aerodynamic environments. In addition, a radiometer design will be matured to measure the radiative component of the reentry heating at two locations on the heat shield. For the back shell, surface thermocouple and pressure port designs will be developed and applied which build upon the heritage of the Space Shuttle Program for instrumentation of reusable surface insulation (RSI) tiles. The quantity and location of the sensors has been determined to balance the needs of the reentry disciplines with the demands of the hardware development, manufacturing and integration. Measurements which provided low relative value and presented significant engineering development effort were, unfortunately, eliminated. The final TPS instrumentation has been optimized to target priority test objectives. The data obtained will serve to provide a better understanding of reentry environments for the Orion capsule design, reduce margins, and potentially reduce TPS mass or provide TPS extensibility for alternative missions.

Boattail Plates With Non-Rectangular Geometries For Reducing Aerodynamic Base Drag Of A Bluff Body In Ground Effect

DOEpatents

Ortega, Jason M.; Sabari, Kambiz

2006-03-07

An apparatus for reducing the aerodynamic base drag of a bluff body having a leading end, a trailing end, a top surface, opposing left and right side surfaces, and a base surface at the trailing end substantially normal to a longitudinal centerline of the bluff body, with the base surface joined (1) to the left side surface at a left trailing edge, (2) to the right side surface at a right trailing edge, and (3) to the top surface at a top trailing edge. The apparatus includes left and right vertical boattail plates which are orthogonally attached to the base surface of the bluff body and inwardly offset from the left and right trailing edges, respectively. This produces left and right vertical channels which generate, in a flowstream substantially parallel to the longitudinal centerline, respective left and right vertically-aligned vortical structures, with the left and right vertical boattail plates each having a plate width defined by a rear edge of the plate spaced from the base surface. Each plate also has a peak plate width at a location between top and bottom ends of the plate corresponding to a peak vortex of the respective vertically-aligned vortical structures.

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.

The Aeroacoustics and Aerodynamics of High-Speed Coanda Devices, Part 1: Conventional Arrangement of Exit Nozzle and Surface

NASA Astrophysics Data System (ADS)

Carpenter, P. W.; Green, P. N.

1997-12-01

The literature on high-speed Coanda flows and its applications is reviewed. The lack of basic information for design engineers is noted. The present paper is based on an investigation of the aeroacoustics and aerodynamics of the high-speed Coanda flow that is formed when a supersonic jet issues from a radial nozzle and adheres to a tulip-shaped body of revolution. Schlieren and other flow visualization techniques together with theoretical methods are used to reveal the various features of this complex flow field. The acoustic characteristics were obtained from measurements with an array of microphones in an anechoic chamber. The emphasis is placed on those features of the aerodynamics and aeroacoustics which may be of general interest.

Aerodynamic preliminary analysis system. Part 2: User's manual and program description

NASA Technical Reports Server (NTRS)

Divan, P.; Dunn, K.; Kojima, J.

1978-01-01

A comprehensive aerodynamic analysis program based on linearized potential theory is described. The solution treats thickness and attitude problems at subsonic and supersonic speeds. Three dimensional configurations with or without jet flaps having multiple nonplanar surfaces of arbitrary planform and open or closed slender bodies or noncircular contour are analyzed. Longitudinal and lateral-directional static and rotary derivative solutions are generated. The analysis is implemented on a time sharing system in conjunction with an input tablet digitizer and an interactive graphics input/output display and editing terminal to maximize its responsiveness to the preliminary analysis problem. Nominal case computation time of 45 CPU seconds on the CDC 175 for a 200 panel simulation indicates the program provides an efficient analysis for systematically performing various aerodynamic configuration tradeoff and evaluation studies.

Results of investigations on a 0.0405 scale model PRR version of the NR-SSV orbiter in the North American Aeronautical Laboratory low speed wind tunnel

NASA Technical Reports Server (NTRS)

Kingsland, R. B.; Vaughn, J. E.; Singellton, R.

1973-01-01

Experimental aerodynamic investigations were conducted in a low speed wind tunnel on a scale model space shuttle vehicle (SSV) orbiter. The purpose of the test was to investigate the longitudinal and lateral-directional aerodynamic characteristics of the space shuttle orbiter. Emphasis was placed on model component, wing-glove, and wing-body fairing effects, as well as elevon, aileron, and rudder control effectiveness. Angles of attack from - 5 deg to + 30 deg and angles of sideslip of - 5 deg, 0 deg, and + 5 deg were tested. Static pressures were recorded on base, fuselage, and wing surfaces. Tufts and talc-kerosene flow visualization techniques were also utilized. The aerodynamic force balance results are presented in plotted and tabular form.

Aerodynamic Leidenfrost effect

NASA Astrophysics Data System (ADS)

Gauthier, Anaïs; Bird, James C.; Clanet, Christophe; Quéré, David

2016-12-01

When deposited on a plate moving quickly enough, any liquid can levitate as it does when it is volatile on a very hot solid (Leidenfrost effect). In the aerodynamic Leidenfrost situation, air gets inserted between the liquid and the moving solid, a situation that we analyze. We observe two types of entrainment. (i) The thickness of the air gap is found to increase with the plate speed, which is interpreted in the Landau-Levich-Derjaguin frame: Air is dynamically dragged along the surface and its thickness results from a balance between capillary and viscous effects. (ii) Air set in motion by the plate exerts a force on the levitating liquid. We discuss the magnitude of this aerodynamic force and show that it can be exploited to control the liquid and even to drive it against gravity.

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.

Flight Test Maneuvers for Efficient Aerodynamic Modeling

NASA Technical Reports Server (NTRS)

Morelli, Eugene A.

2011-01-01

Novel flight test maneuvers for efficient aerodynamic modeling were developed and demonstrated in flight. Orthogonal optimized multi-sine inputs were applied to aircraft control surfaces to excite aircraft dynamic response in all six degrees of freedom simultaneously while keeping the aircraft close to chosen reference flight conditions. Each maneuver was designed for a specific modeling task that cannot be adequately or efficiently accomplished using conventional flight test maneuvers. All of the new maneuvers were first described and explained, then demonstrated on a subscale jet transport aircraft in flight. Real-time and post-flight modeling results obtained using equation-error parameter estimation in the frequency domain were used to show the effectiveness and efficiency of the new maneuvers, as well as the quality of the aerodynamic models that can be identified from the resultant flight data.

Aircraft geometry verification with enhanced computer generated displays

NASA Technical Reports Server (NTRS)

Cozzolongo, J. V.

1982-01-01

A method for visual verification of aerodynamic geometries using computer generated, color shaded images is described. The mathematical models representing aircraft geometries are created for use in theoretical aerodynamic analyses and in computer aided manufacturing. The aerodynamic shapes are defined using parametric bi-cubic splined patches. This mathematical representation is then used as input to an algorithm that generates a color shaded image of the geometry. A discussion of the techniques used in the mathematical representation of the geometry and in the rendering of the color shaded display is presented. The results include examples of color shaded displays, which are contrasted with wire frame type displays. The examples also show the use of mapped surface pressures in terms of color shaded images of V/STOL fighter/attack aircraft and advanced turboprop aircraft.

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.

Surface Microwave and Surface Transversal Pulsed-Periodic Discharges in Supersonic Flow

DTIC Science & Technology

2004-03-01

plasmas of different types of gas discharges near the surface of Aerodynamic models and in the boundary layers. Also, the contractor will develop modes...regions near the surface. The following experimental work will be done in supersonic air flow (Mɚ) at pressures between 1 and 200 Torr: a...198 CHAPTER IX NUMERICAL CALCULATION OF CHARACTERISTICS OF SUPERSONIC FLOW NEAR A FLAT PLATE WITH MICROWAVE DISCHARGE ON ITS SURFACE

Contributions of the NASA Langley Research Center to the DARPA/AFRL/NASA/ Northrop Grumman Smart Wing Program

NASA Technical Reports Server (NTRS)

Florance, Jennifer P.; Burner, Alpheus W.; Fleming, Gary A.; Martin, Christopher A.

2003-01-01

An overview of the contributions of the NASA Langley Research Center (LaRC) to the DARPA/AFRL/NASA/ Northrop Grumman Corporation (NGC) Smart Wing program is presented. The overall objective of the Smart Wing program was to develop smart** technologies and demonstrate near-flight-scale actuation systems to improve the aerodynamic performance of military aircraft. NASA LaRC s roles were to provide technical guidance, wind-tunnel testing time and support, and Computational Fluid Dynamics (CFD) analyses. The program was divided into two phases, with each phase having two wind-tunnel entries in the Langley Transonic Dynamics Tunnel (TDT). This paper focuses on the fourth and final wind-tunnel test: Phase 2, Test 2. During this test, a model based on the NGC Unmanned Combat Air Vehicle (UCAV) concept was tested at Mach numbers up to 0.8 and dynamic pressures up to 150 psf to determine the aerodynamic performance benefits that could be achieved using hingeless, smoothly-contoured control surfaces actuated with smart materials technologies. The UCAV-based model was a 30% geometric scale, full-span, sting-mounted model with the smart control surfaces on the starboard wing and conventional, hinged control surfaces on the port wing. Two LaRC-developed instrumentation systems were used during the test to externally measure the shapes of the smart control surface and quantify the effects of aerodynamic loading on the deflections: Videogrammetric Model Deformation (VMD) and Projection Moire Interferometry (PMI). VMD is an optical technique that uses single-camera photogrammetric tracking of discrete targets to determine deflections at specific points. PMI provides spatially continuous measurements of model deformation by computationally analyzing images of a grid projected onto the model surface. Both the VMD and PMI measurements served well to validate the use of on-board (internal) rotary potentiometers to measure the smart control surface deflection angles. Prior to the final entry, NASA LaRC also performed three-dimensional unstructured Navier Stokes CFD analyses in an attempt to predict the potential aerodynamic impact of the smart control surface on overall model forces and moments. Eight different control surface shapes were selected for study at Mach = 0.6, Reynolds number = 3.25 x 10(exp 6), and + 2 deg., 3 deg., 8 deg., and 10 deg.model angles-of-attack. For the baseline, undeflected control surface geometry, the CFD predictions and wind-tunnel results matched well. The agreement was not as good for the more complex aero-loaded control surface shapes, though, because of the inability to accurately predict those shapes. Despite these results, the NASA CFD study served as an important step in studying advanced control effectors.

Aerodynamic surface stress intermittency and conditionally averaged turbulence statistics

NASA Astrophysics Data System (ADS)

Anderson, William; Lanigan, David

2015-11-01

Aeolian erosion is induced by aerodynamic stress imposed by atmospheric winds. Erosion models prescribe that sediment flux, Q, scales with aerodynamic stress raised to exponent, n, where n > 1 . Since stress (in fully rough, inertia-dominated flows) scales with incoming velocity squared, u2, it follows that q ~u2n (where u is some relevant component of the flow). Thus, even small (turbulent) deviations of u from its time-mean may be important for aeolian activity. This rationale is augmented given that surface layer turbulence exhibits maximum Reynolds stresses in the fluid immediately above the landscape. To illustrate the importance of stress intermittency, we have used conditional averaging predicated on stress during large-eddy simulation of atmospheric boundary layer flow over an arid, bare landscape. Conditional averaging provides an ensemble-mean visualization of flow structures responsible for erosion `events'. Preliminary evidence indicates that surface stress peaks are associated with the passage of inclined, high-momentum regions flanked by adjacent low-momentum regions. We characterize geometric attributes of such structures and explore streamwise and vertical vorticity distribution within the conditionally averaged flow field. This work was supported by the National Sci. Foundation, Phys. and Dynamic Meteorology Program (PM: Drs. N. Anderson, C. Lu, and E. Bensman) under Grant # 1500224. Computational resources were provided by the Texas Adv. Comp. Center at the Univ. of Texas.

Space shuttle propulsion parameter estimation using optimal estimation techniques

NASA Technical Reports Server (NTRS)

1983-01-01

The first twelve system state variables are presented with the necessary mathematical developments for incorporating them into the filter/smoother algorithm. Other state variables, i.e., aerodynamic coefficients can be easily incorporated into the estimation algorithm, representing uncertain parameters, but for initial checkout purposes are treated as known quantities. An approach for incorporating the NASA propulsion predictive model results into the optimal estimation algorithm was identified. This approach utilizes numerical derivatives and nominal predictions within the algorithm with global iterations of the algorithm. The iterative process is terminated when the quality of the estimates provided no longer significantly improves.

In-flight investigation of shuttle tile pressure orifice installations

NASA Technical Reports Server (NTRS)

Moes, Timothy R.; Meyer, Robert R., Jr.

1990-01-01

To determine shuttle orbiter wing loads during ascent, wing load instrumentation was added to Columbia (OV-102). This instrumentation included strain gages and pressure orifices on the wing. The loads derived from wing pressure measurements taken during STS 61-C did not agree with those derived from strain gage measurements or with the loads predicted from the aerodynamic database. Anomalies in the surface immediately surrounding the pressure orifices in the thermal protection system (TPS) tiles were one possible cause of errors in the loads derived from wing pressure measurements. These surface anomalies were caused by a ceramic filler material which was installed around the pressure tubing. The filler material allowed slight movement of the TPS tile and pressure tube as the airframe flexed and bent under aerodynamic loads during ascent and descent. Postflight inspection revealed that this filler material had protruded from or receeded beneath the surface, causing the orifice to lose its flushness. Flight tests were conducted at NASA Ames Research Center Dryden Flight Research Facility to determine the effects of any anomaly in surface flushness of the orifice installation on the measured pressures at Mach numbers between 0.6 and 1.4. An F-104 aircraft with a flight test fixture mounted beneath the fuselage was used for these flights. Surface flushness anomalies typical of those on the orbiter after flight (STA 61-C) were tested. Also, cases with excessive protrusion and recession of the filler material were tested. This report shows that the anomalies in STS 61-C orifice installations adversely affected the pressure measurements. But the magnitude of the affect was not great enough to account for the discrepancies with the strain gage measurements and the aerodynamic predictions.

Uncertainty Analysis of Instrument Calibration and Application

NASA Technical Reports Server (NTRS)

Tripp, John S.; Tcheng, Ping

1999-01-01

Experimental aerodynamic researchers require estimated precision and bias uncertainties of measured physical quantities, typically at 95 percent confidence levels. Uncertainties of final computed aerodynamic parameters are obtained by propagation of individual measurement uncertainties through the defining functional expressions. In this paper, rigorous mathematical techniques are extended to determine precision and bias uncertainties of any instrument-sensor system. Through this analysis, instrument uncertainties determined through calibration are now expressed as functions of the corresponding measurement for linear and nonlinear univariate and multivariate processes. Treatment of correlated measurement precision error is developed. During laboratory calibration, calibration standard uncertainties are assumed to be an order of magnitude less than those of the instrument being calibrated. Often calibration standards do not satisfy this assumption. This paper applies rigorous statistical methods for inclusion of calibration standard uncertainty and covariance due to the order of their application. The effects of mathematical modeling error on calibration bias uncertainty are quantified. The effects of experimental design on uncertainty are analyzed. The importance of replication is emphasized, techniques for estimation of both bias and precision uncertainties using replication are developed. Statistical tests for stationarity of calibration parameters over time are obtained.

Unified aeroacoustics analysis for high speed turboprop aerodynamics and noise. Volume 3: Application of theory for blade loading, wakes, noise, and wing shielding

NASA Technical Reports Server (NTRS)

Hanson, D. B.; Mccolgan, C. J.; Ladden, R. M.; Klatte, R. J.

1991-01-01

Results of the program for the generation of a computer prediction code for noise of advanced single rotation, turboprops (prop-fans) such as the SR3 model are presented. The code is based on a linearized theory developed at Hamilton Standard in which aerodynamics and acoustics are treated as a unified process. Both steady and unsteady blade loading are treated. Capabilities include prediction of steady airload distributions and associated aerodynamic performance, unsteady blade pressure response to gust interaction or blade vibration, noise fields associated with thickness and steady and unsteady loading, and wake velocity fields associated with steady loading. The code was developed on the Hamilton Standard IBM computer and has now been installed on the Cray XMP at NASA-Lewis. The work had its genesis in the frequency domain acoustic theory developed at Hamilton Standard in the late 1970s. It was found that the method used for near field noise predictions could be adapted as a lifting surface theory for aerodynamic work via the pressure potential technique that was used for both wings and ducted turbomachinery. In the first realization of the theory for propellers, the blade loading was represented in a quasi-vortex lattice form. This was upgraded to true lifting surface loading. Originally, it was believed that a purely linear approach for both aerodynamics and noise would be adequate. However, two sources of nonlinearity in the steady aerodynamics became apparent and were found to be a significant factor at takeoff conditions. The first is related to the fact that the steady axial induced velocity may be of the same order of magnitude as the flight speed and the second is the formation of leading edge vortices which increases lift and redistribute loading. Discovery and properties of prop-fan leading edge vortices were reported in two papers. The Unified AeroAcoustic Program (UAAP) capabilites are demonstrated and the theory verified by comparison with the predictions with data from tests at NASA-Lewis. Steady aerodyanmic performance, unsteady blade loading, wakes, noise, and wing and boundary layer shielding are examined.

Detecting surface roughness effects on the atmospheric boundary layer via AIRSAR data: A field experiment in Death Valley, California

NASA Technical Reports Server (NTRS)

Blumberg, Dan G.; Greeley, Ronald

1992-01-01

The part of the troposphere influenced by the surface of the earth is termed the atmospheric boundary layer. Flow within this layer is influenced by the roughness of the surface; rougher surfaces induce more turbulence than smoother surfaces and, hence, higher atmospheric transfer rates across the surface. Roughness elements also shield erodible particles, thus decreasing the transport of windblown particles. Therefore, the aerodynamic roughness length (z(sub 0)) is an important parameter in aeolian and atmospheric boundary layer processes as it describes the aerodynamic properties of the underlying surface. z(sub 0) is assumed to be independent of wind velocity or height, and dependent only on the surface topography. It is determined using in situ measurements of the wind speed distribution as a function of height. For dry, unvegetated soils the intensity of the radar backscatter (sigma(sup 0)) is affected primarily by surface roughness at a scale comparable with the radar wavelength. Thus, both wind and radar respond to surface roughness variations on a scale of a few meters or less. Greeley showed the existence of a correlation between z(sub 0) and sigma(sup 0). This correlation was based on measurements over lava flows, alluvial fans, and playas in the southwest deserts of the United States. It is shown that the two parameters behave similarly also when there are small changes over a relatively homogeneous surface.

Characteristics of the Martian atmosphere surface layer

NASA Technical Reports Server (NTRS)

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

1990-01-01

Elements of various terrestrial boundary layer models are extended to Mars in order to estimate sensible heat, latent heat, and momentum fluxes within the Martian atmospheric surface ('constant flux') layer. The atmospheric surface layer consists of an interfacial sublayer immediately adjacent to the ground and an overlying fully turbulent surface sublayer where wind-shear production of turbulence dominates buoyancy production. Within the interfacial sublayer, sensible and latent heat are transported by non-steady molecular diffusion into small-scale eddies which intermittently burst through this zone. Both the thickness of the interfacial sublayer and the characteristics of the turbulent eddies penetrating through it depend on whether airflow is aerodynamically smooth or aerodynamically rough, as determined by the Roughness Reynold's number. Within the overlying surface sublayer, similarity theory can be used to express the mean vertical windspeed, temperature, and water vapor profiles in terms of a single parameter, the Monin-Obukhov stability parameter. To estimate the molecular viscosity and thermal conductivity of a CO2-H2O gas mixture under Martian conditions, parameterizations were developed using data from the TPRC Data Series and the first-order Chapman-Cowling expressions; the required collision integrals were approximated using the Lenard-Jones potential. Parameterizations for specific heat and binary diffusivity were also determined. The Brutsart model for sensible and latent heat transport within the interfacial sublayer for both aerodynamically smooth and rough airflow was experimentally tested under similar conditions, validating its application to Martian conditions. For the surface sublayer, the definition of the Monin-Obukhov length was modified to properly account for the buoyancy forces arising from water vapor gradients in the Martian atmospheric boundary layer. It was found that under most Martian conditions, the interfacial and surface sublayers offer roughly comparable resistance to sensible heat and water vapor transport and are thus both important in determining the associated fluxes.

Development of Variable Camber Continuous Trailing Edge Flap for Performance Adaptive Aeroelastic Wing

NASA Technical Reports Server (NTRS)

Nguyen, Nhan; Kaul, Upender; Lebofsky, Sonia; Ting, Eric; Chaparro, Daniel; Urnes, James

2015-01-01

This paper summarizes the recent development of an adaptive aeroelastic wing shaping control technology called variable camber continuous trailing edge flap (VCCTEF). As wing flexibility increases, aeroelastic interactions with aerodynamic forces and moments become an increasingly important consideration in aircraft design and aerodynamic performance. Furthermore, aeroelastic interactions with flight dynamics can result in issues with vehicle stability and control. The initial VCCTEF concept was developed in 2010 by NASA under a NASA Innovation Fund study entitled "Elastically Shaped Future Air Vehicle Concept," which showed that highly flexible wing aerodynamic surfaces can be elastically shaped in-flight by active control of wing twist and bending deflection in order to optimize the spanwise lift distribution for drag reduction. A collaboration between NASA and Boeing Research & Technology was subsequently funded by NASA from 2012 to 2014 to further develop the VCCTEF concept. This paper summarizes some of the key research areas conducted by NASA during the collaboration with Boeing Research and Technology. These research areas include VCCTEF design concepts, aerodynamic analysis of VCCTEF camber shapes, aerodynamic optimization of lift distribution for drag minimization, wind tunnel test results for cruise and high-lift configurations, flutter analysis and suppression control of flexible wing aircraft, and multi-objective flight control for adaptive aeroelastic wing shaping control.

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.

Predicting aerodynamic characteristics of vortical flows on three-dimensional configurations using a surface-singularity panel method

NASA Technical Reports Server (NTRS)

Maskew, B.

1983-01-01

A general low-order surface-singularity panel method is used to predict the aerodynamic characteristics of a problem where a wing-tip vortex from one wing closely interacts with an aft mounted wing in a low Reynolds Number flow; i.e., 125,000. Nonlinear effects due to wake roll-up and the influence of the wings on the vortex path are included in the calculation by using a coupled iterative wake relaxation scheme. The interaction also affects the wing pressures and boundary layer characteristics: these effects are also considered using coupled integral boundary layer codes and preliminary calculations using free vortex sheet separation modelling are included. Calculated results are compared with water tunnel experimental data with generally remarkably good agreement.

Suction and Blowing Flow Control on Airfoil for Drag Reduction in Subsonic Flow

NASA Astrophysics Data System (ADS)

Baljit, S. S.; Saad, M. R.; Nasib, A. Z.; Sani, A.; Rahman, M. R. A.; Idris, A. C.

2017-10-01

Lift force is produced from a pressure difference between the pressures acting in upper and lower surfaces. Therefore, flow becomes detached from the surface of the airfoil at separation point and form vortices. These vortices affect the aerodynamic performance of the airfoil in term of lift and drag coefficient. Therefore, this study is investigating the effect of suction and jet blowing in boundary layer separation control on NACA 0012 airfoil in a subsonic wind tunnel. The experiment examined both methods at the position of 25% of the chord-length of the airfoil at Reynolds number 1.2 × 105. The findings show that suction and jet blowing affect the aerodynamic performance of NACA 0012 airfoil and can be an effective means for boundary layer separation control in subsonic flow.

Elevated-Temperature Tests Under Static and Aerodynamic Conditions on Honeycomb-Core Sandwich Panels

NASA Technical Reports Server (NTRS)

Groen, Joseph M.; Johnson, Aldie E., Jr.

1959-01-01

Stainless-steel honeycomb-core sandwich panels which differed primarily in skin thicknesses were tested at elevated temperatures under static and aerodynamic conditions. The results of these tests were evaluated to determine the insulating effectiveness and structural integrity of the panels. The static radiant-heating tests were performed in front of a quartz-tube radiant heater at panel skin temperatures up to 1,5000 F. The aerodynamic tests were made in a Mach 1.4 heated blowdown wind tunnel. The tunnel temperature was augmented by additional heat supplied by a radiant heater which raised the panel surface temperature above 8000 F during air flow. Static radiant-heating tests of 2 minutes duration showed that all the panels protected the load-carrying structure about equally well. Thin-skin panels showed an advantage for this short-time test over thick-skin panels from a standpoint of weight against insulation. Permanent inelastic strains in the form of local buckles over each cell of the honeycomb core caused an increase in surface roughness. During the aero- dynamic tests all of the panels survived with little or no damage, and panel flutter did not occur.

Beneficial aerodynamic effect of wing scales on the climbing flight of butterflies.

PubMed

Slegers, Nathan; Heilman, Michael; Cranford, Jacob; Lang, Amy; Yoder, John; Habegger, Maria Laura

2017-01-30

It is hypothesized that butterfly wing scale geometry and surface patterning may function to improve aerodynamic efficiency. In order to investigate this hypothesis, a method to measure butterfly flapping kinematics optically over long uninhibited flapping sequences was developed. Statistical results for the climbing flight flapping kinematics of 11 butterflies, based on a total of 236 individual flights, both with and without their wing scales, are presented. Results show, that for each of the 11 butterflies, the mean climbing efficiency decreased after scales were removed. Data was reduced to a single set of differences of climbing efficiency using are paired t-test. Results show a mean decrease in climbing efficiency of 32.2% occurred with a 95% confidence interval of 45.6%-18.8%. Similar analysis showed that the flapping amplitude decreased by 7% while the flapping frequency did not show a significant difference. Results provide strong evidence that butterfly wing scale geometry and surface patterning improve butterfly climbing efficiency. The authors hypothesize that the wing scale's effect in measured climbing efficiency may be due to an improved aerodynamic efficiency of the butterfly and could similarly be used on flapping wing micro air vehicles to potentially achieve similar gains in efficiency.

2-D and 3-D oscillating wing aerodynamics for a range of angles of attack including stall

NASA Technical Reports Server (NTRS)

Piziali, R. A.

1994-01-01

A comprehensive experimental investigation of the pressure distribution over a semispan wing undergoing pitching motions representative of a helicopter rotor blade was conducted. Testing the wing in the nonrotating condition isolates the three-dimensional (3-D) blade aerodynamic and dynamic stall characteristics from the complications of the rotor blade environment. The test has generated a very complete, detailed, and accurate body of data. These data include static and dynamic pressure distributions, surface flow visualizations, two-dimensional (2-D) airfoil data from the same model and installation, and important supporting blockage and wall pressure distributions. This body of data is sufficiently comprehensive and accurate that it can be used for the validation of rotor blade aerodynamic models over a broad range of the important parameters including 3-D dynamic stall. This data report presents all the cycle-averaged lift, drag, and pitching moment coefficient data versus angle of attack obtained from the instantaneous pressure data for the 3-D wing and the 2-D airfoil. Also presented are examples of the following: cycle-to-cycle variations occurring for incipient or lightly stalled conditions; 3-D surface flow visualizations; supporting blockage and wall pressure distributions; and underlying detailed pressure results.

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.

Application of a quasi-3D inviscid flow and boundary layer analysis to the hub-shroud contouring of a radial turbine

NASA Technical Reports Server (NTRS)

Civinskas, K.; Povinelli, L. A.

1984-01-01

Application of a quasi-3D approach to the aerodynamic analysis of several radial turbine configurations is described. The objective was to improve the rotor aerodynamic characteristics by hub-shroud contouring. The approach relies on available 2D inviscid methods coupled with boundary layer analysis to calculate profile, mixing, and endwall losses. Windage, tip clearance, incidence, and secondary flow losses are estimated from correlations. To eliminate separation along the hub and blade suction surfaces of a baseline rotor, the analysis was also applied to three alternate hub-shroud geometries. Emphasis was on elimination of an inducer velocity overshoot as well as increasing hub velocities. While separation was never eliminated, the extent of the separated area was progressively reduced. Results are presented in terms of mid-channel and blade surface velocities; kinetic energy loss coefficients; and efficiency. The calculation demonstrates a first step for a systematic approach to radial turbine design that can be used to identify and control aerodynamic characteristics that ultimately determine heat transfer and component life. Experimentation will be required to assess the extent to which flow and boundary layer behavior were predicted correctly.

Application of a quasi-3D inviscid flow and boundary layer analysis to the hub-shroud contouring of a radial turbine

NASA Technical Reports Server (NTRS)

Civinskas, K. C.; Povinelli, L. A.

1984-01-01

Application of a quasi-3D approach to the aerodynamic analysis of several radial turbine configurations is described. The objective was to improve the rotor aerodynamic characteristics by hub-shroud contouring. The approach relies on available 2D inviscid methods coupled with boundary layer analysis to calculate profile, mixing, and endwall losses. Windage, tip clearance, incidence, and secondary flow losses are estimated from correlations. To eliminate separation along the hub and blade suction surfaces of a baseline rotor, the analysis was also applied to three alternate hub-shroud geometries. Emphasis was on elimination an inducer velocity overshoot as well as increasing hub velocities. While separation was never eliminated, the extent of the separated area was progressively reduced. Results are presented in terms of mid-channel and blade surface velocities; kinetic energy loss coefficients; and efficiency. The calculation demonstrates a first step for a systematic approach to radial turbine design that can be used to identify and control aerodynamic characteristics that ultimately determine heat transfer and component life. Experimentation will be required to assess the extent to which flow and boundary layer behavior were predicted correctly.

Control of vortex on a non-slender delta wing by a nanosecond pulse surface dielectric barrier discharge

NASA Astrophysics Data System (ADS)

Zhao, Guang-yin; Li, Ying-hong; Liang, Hua; Han, Meng-hu; Hua, Wei-zhuo

2015-01-01

Wind tunnel experiments are conducted for improving the aerodynamic performance of delta wing using a leading-edge pulsed nanosecond dielectric barrier discharge (NS-DBD). The whole effects of pulsed NS-DBD on the aerodynamic performance of the delta wing are studied by balanced force measurements. Pressure measurements and particle image velocimetry (PIV) measurements are conducted to investigate the formation of leading-edge vortices affected by the pulsed NS-DBD, compared to completely stalled flow without actuation. Various pulsed actuation frequencies of the plasma actuator are examined with the freestream velocity up to 50 m/s. Stall has been delayed substantially and significant shifts in the aerodynamic forces can be achieved at the post-stall regions when the actuator works at the optimum reduced frequency of F + = 2. The upper surface pressure measurements show that the largest change of static pressure occurs at the forward part of the wing at the stall region. The time-averaged flow pattern obtained from the PIV measurement shows that flow reattachment is promoted with excitation, and a vortex flow pattern develops. The time-averaged locations of the secondary separation line and the center of the vortical region both move outboard with excitation.

Actuator and aerodynamic modeling for high-angle-of-attack aeroservoelasticity

NASA Technical Reports Server (NTRS)

Brenner, Martin J.

1993-01-01

Accurate prediction of airframe/actuation coupling is required by the imposing demands of modern flight control systems. In particular, for agility enhancement at high angle of attack and low dynamic pressure, structural integration characteristics such as hinge moments, effective actuator stiffness, and airframe/control surface damping can have a significant effect on stability predictions. Actuator responses are customarily represented with low-order transfer functions matched to actuator test data, and control surface stiffness is often modeled as a linear spring. The inclusion of the physical properties of actuation and its installation on the airframe is therefore addressed in this paper using detailed actuator models which consider the physical, electrical, and mechanical elements of actuation. The aeroservoelastic analysis procedure is described in which the actuators are modeled as detailed high-order transfer functions and as approximate low-order transfer functions. The impacts of unsteady aerodynamic modeling on aeroservoelastic stability are also investigated in this paper by varying the order of approximation, or number of aerodynamic lag states, in the analysis. Test data from a thrust-vectoring configuration of an F/A-18 aircraft are compared to predictions to determine the effects on accuracy as a function of modeling complexity.

Actuator and aerodynamic modeling for high-angle-of-attack aeroservoelasticity

NASA Technical Reports Server (NTRS)

Brenner, Martin J.

1993-01-01

Accurate prediction of airframe/actuation coupling is required by the imposing demands of modern flight control systems. In particular, for agility enhancement at high angle of attack and low dynamic pressure, structural integration characteristics such as hinge moments, effective actuator stiffness, and airframe/control surface damping can have a significant effect on stability predictions. Actuator responses are customarily represented with low-order transfer functions matched to actuator test data, and control surface stiffness is often modeled as a linear spring. The inclusion of the physical properties of actuation and its installation on the airframe is therefore addressed using detailed actuator models which consider the physical, electrical, and mechanical elements of actuation. The aeroservoelastic analysis procedure is described in which the actuators are modeled as detailed high-order transfer functions and as approximate low-order transfer functions. The impacts of unsteady aerodynamic modeling on aeroservoelastic stability are also investigated by varying the order of approximation, or number of aerodynamic lag states, in the analysis. Test data from a thrust-vectoring configuration of an F/A-l8 aircraft are compared to predictions to determine the effects on accuracy as a function of modeling complexity.

Introduction to Generalized Functions with Applications in Aerodynamics and Aeroacoustics

NASA Technical Reports Server (NTRS)

Farassat, F.

1994-01-01

Generalized functions have many applications in science and engineering. One useful aspect is that discontinuous functions can be handled as easily as continuous or differentiable functions and provide a powerful tool in formulating and solving many problems of aerodynamics and acoustics. Furthermore, generalized function theory elucidates and unifies many ad hoc mathematical approaches used by engineers and scientists. We define generalized functions as continuous linear functionals on the space of infinitely differentiable functions with compact support, then introduce the concept of generalized differentiation. Generalized differentiation is the most important concept in generalized function theory and the applications we present utilize mainly this concept. First, some results of classical analysis, are derived with the generalized function theory. Other applications of the generalized function theory in aerodynamics discussed here are the derivations of general transport theorems for deriving governing equations of fluid mechanics, the interpretation of the finite part of divergent integrals, the derivation of the Oswatitsch integral equation of transonic flow, and the analysis of velocity field discontinuities as sources of vorticity. Applications in aeroacoustics include the derivation of the Kirchhoff formula for moving surfaces, the noise from moving surfaces, and shock noise source strength based on the Ffowcs Williams-Hawkings equation.

Aerodynamic parameters from distributed heterogeneous CNT hair sensors with a feedforward neural network.

PubMed

Magar, Kaman Thapa; Reich, Gregory W; Kondash, Corey; Slinker, Keith; Pankonien, Alexander M; Baur, Jeffery W; Smyers, Brian

2016-11-10

Distributed arrays of artificial hair sensors have bio-like sensing capabilities to obtain spatial and temporal surface flow information which is an important aspect of an effective fly-by-feel system. The spatiotemporal surface flow measurement enables further exploration of additional flow features such as flow stagnation, separation, and reattachment points. Due to their inherent robustness and fault tolerant capability, distributed arrays of hair sensors are well equipped to assess the aerodynamic and flow states in adverse conditions. In this paper, a local flow measurement from an array of artificial hair sensors in a wind tunnel experiment is used with a feedforward artificial neural network to predict aerodynamic parameters such as lift coefficient, moment coefficient, free-stream velocity, and angle of attack on an airfoil. We find the prediction error within 6% and 10% for lift and moment coefficients. The error for free-stream velocity and angle of attack were within 0.12 mph and 0.37 degrees. Knowledge of these parameters are key to finding the real time forces and moments which paves the way for effective control design to increase flight agility, stability, and maneuverability.

Simulation and Flight Evaluation of a Parameter Estimation Input Design Method for Hybrid-Wing-Body Aircraft

NASA Technical Reports Server (NTRS)

Taylor, Brian R.; Ratnayake, Nalin A.

2010-01-01

As part of an effort to improve emissions, noise, and performance of next generation aircraft, it is expected that future aircraft will make use of distributed, multi-objective control effectors in a closed-loop flight control system. Correlation challenges associated with parameter estimation will arise with this expected aircraft configuration. Research presented in this paper focuses on addressing the correlation problem with an appropriate input design technique and validating this technique through simulation and flight test of the X-48B aircraft. The X-48B aircraft is an 8.5 percent-scale hybrid wing body aircraft demonstrator designed by The Boeing Company (Chicago, Illinois, USA), built by Cranfield Aerospace Limited (Cranfield, Bedford, United Kingdom) and flight tested at the National Aeronautics and Space Administration Dryden Flight Research Center (Edwards, California, USA). Based on data from flight test maneuvers performed at Dryden Flight Research Center, aerodynamic parameter estimation was performed using linear regression and output error techniques. An input design technique that uses temporal separation for de-correlation of control surfaces is proposed, and simulation and flight test results are compared with the aerodynamic database. This paper will present a method to determine individual control surface aerodynamic derivatives.

Screening of the aerodynamic and biophysical properties of barley malt

NASA Astrophysics Data System (ADS)

Ghodsvali, Alireza; Farzaneh, Vahid; Bakhshabadi, Hamid; Zare, Zahra; Karami, Zahra; Mokhtarian, Mohsen; Carvalho, Isabel. S.

2016-10-01

An understanding of the aerodynamic and biophysical properties of barley malt is necessary for the appropriate design of equipment for the handling, shipping, dehydration, grading, sorting and warehousing of this strategic crop. Malting is a complex biotechnological process that includes steeping; germination and finally, the dehydration of cereal grains under controlled temperature and humidity conditions. In this investigation, the biophysical properties of barley malt were predicted using two models of artificial neural networks as well as response surface methodology. Stepping time and germination time were selected as the independent variables and 1 000 kernel weight, kernel density and terminal velocity were selected as the dependent variables (responses). The obtained outcomes showed that the artificial neural network model, with a logarithmic sigmoid activation function, presents more precise results than the response surface model in the prediction of the aerodynamic and biophysical properties of produced barley malt. This model presented the best result with 8 nodes in the hidden layer and significant correlation coefficient values of 0.783, 0.767 and 0.991 were obtained for responses one thousand kernel weight, kernel density, and terminal velocity, respectively. The outcomes indicated that this novel technique could be successfully applied in quantitative and qualitative monitoring within the malting process.

Reference H Cycle 3 Stability, Control, and Flying Qualities Batch Assessments

NASA Technical Reports Server (NTRS)

Henderson, Dennis K.

1999-01-01

This work is an update of the assessment completed in February of 1996, when a preliminary assessment report was issued for the Cycle 2B simulation model. The primary purpose of the final assessment was to re-evaluate each assessment against the flight control system (FCS) requirements document using the updated model. Only a limited number of final assessments were completed due to the close proximity of the release of the Langley model and the assessment deliverable date. The assessment used the nonlinear Cycle 3 simulation model because it combines nonlinear aeroelastic (quasi-static) aerodynamic with hinge moment and rate limited control surface deflections. Both Configuration Aerodynamics (Task 32) and Flight Controls (Task 36) were funded in 1996 to conduct the final stability and control assessments of the unaugmented Reference H configuration in FY96. Because the two tasks had similar output requirements, the work was divided such that Flight Controls would be responsible for the implementation and checkout of the simulation model and Configuration Aerodynamics for writing Madab "script' files, conducting the batch assessments and writing the assessment report. Additionally, Flight Controls was to investigate control surface allocations schemes different from the baseline Reference H in an effort to fulfill flying qualities criteria.

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.

Estimating the aerodynamic roughness of debris covered glacier ice

NASA Astrophysics Data System (ADS)

Quincey, Duncan; Smith, Mark; Rounce, David; Ross, Andrew; King, Owen; Watson, Scott

2017-04-01

Aerodynamic roughness length (z0), the height above the ground surface at which the extrapolated horizontal wind velocity profile drops to zero, is one of the most poorly parameterised elements of the glacier surface energy balance equation. Microtopographic methods for estimating z0 are becoming increasingly well used, but are rarely validated against independent measures and are yet to be comprehensively analysed for scale or data resolution dependency. Here, we present the results of a field investigation conducted on the debris covered Khumbu Glacier during the post-monsoon season of 2015. We focus on two sites. The first is characterised by gravels and cobbles supported by a fine sandy matrix. The second comprises cobbles and boulders separated by voids. Vertical profiles of wind speed measured over both sites enable us to derive measurements of aerodynamic roughness that reflect their observed surface characteristics (0.0184 m vs 0.0243 m). z0 at the second site also varied through time following snowfall (0.0055 m) and during its subsequent melt (0.0129 m), showing the importance of fine resolution topography for near-surface airflow. We conducted Structure from Motion Multi-View Stereo (SfM-MVS) surveys across each patch and calculated z0 using three microtopographic methods. The fully three-dimensional cloud-based approach is shown to be most stable across different scales and these z0 values are most correct in relative order when compared to the wind tower data. Popular profile-based methods perform less well providing highly variable values across different scales and when using data of differing resolution.

Effects of Chemistry on Blunt-Body Wake Structure

NASA Technical Reports Server (NTRS)

Dogra, Virendra K.; Moss, James N.; Wilmoth, Richard G.; Taylor, Jeff C.; Hassan, H. A.

1995-01-01

Results of a numerical study are presented for hypersonic low-density flow about a 70-deg blunt cone using direct simulation Monte Carlo (DSMC) and Navier-Stokes calculations. Particular emphasis is given to the effects of chemistry on the near-wake structure and on the surface quantities and the comparison of the DSMC results with the Navier-Stokes calculations. The flow conditions simulated are those experienced by a space vehicle at an altitude of 85 km and a velocity of 7 km/s during Earth entry. A steady vortex forms in the near wake for these freestream conditions for both chemically reactive and nonreactive air gas models. The size (axial length) of the vortex for the reactive air calculations is 25% larger than that of the nonreactive air calculations. The forebody surface quantities are less sensitive to the chemistry than the base surface quantities. The presence of the afterbody has no effect on the forebody flow structure or the surface quantities. The comparisons of DSMC and Navier-Stokes calculations show good agreement for the wake structure and the forebody surface quantities.

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.

Subsonic wind-tunnel measurements of a slender wing-body configuration employing a vortex flap

NASA Technical Reports Server (NTRS)

Frink, Neal T.

1987-01-01

A wind tunnel study at Mach 0.4 was conducted for a slender wing-body configuration with a leading edge vortex flap of curved planform that is deflectable about a 74 degree swept hinge line. The basic data consist of a unique combination of longitudinal aerodynamic, surface pressure, and vortex flap hinge-moment measurements on a common model. The longitudinal aerodynamic, pressure and hinge-moment data are presented without analysis in tabular format. Plots of the tabulated pressure data are also given.

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.

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.

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.

Aerodynamic preliminary analysis system. Part 1: Theory. [linearized potential theory

NASA Technical Reports Server (NTRS)

Bonner, E.; Clever, W.; Dunn, K.

1978-01-01

A comprehensive aerodynamic analysis program based on linearized potential theory is described. The solution treats thickness and attitude problems at subsonic and supersonic speeds. Three dimensional configurations with or without jet flaps having multiple non-planar surfaces of arbitrary planform and open or closed slender bodies of non-circular contour may be analyzed. Longitudinal and lateral-directional static and rotary derivative solutions may be generated. The analysis was implemented on a time sharing system in conjunction with an input tablet digitizer and an interactive graphics input/output display and editing terminal to maximize its responsiveness to the preliminary analysis problem. Nominal case computation time of 45 CPU seconds on the CDC 175 for a 200 panel simulation indicates the program provides an efficient analysis for systematically performing various aerodynamic configuration tradeoff and evaluation studies.

The effectiveness of vane-aileron excitation in the experimental determination of flutter speed by parameter identification

NASA Technical Reports Server (NTRS)

Nissim, Eli

1990-01-01

The effectiveness of aerodynamic excitation is evaluated analytically in conjunction with the experimental determination of flutter dynamic pressure by parameter identification. Existing control surfaces were used, with an additional vane located at the wingtip. The equations leading to the identification of the equations of motion were reformulated to accommodate excitation forces of aerodynamic origin. The aerodynamic coefficients of the excitation forces do not need to be known since they are determined by the identification procedure. The 12 degree-of-freedom numerical example treated in this work revealed the best wingtip vane locations, and demonstrated the effectiveness of the aileron-vane excitation system. Results from simulated data gathered at much lower dynamic pressures (approximately half the value of flutter dynamic pressure) predicted flutter dynamic pressures with 2-percent errors.

Unified Formulation of the Aeroelasticity of Swept Lifting Surfaces

NASA Technical Reports Server (NTRS)

Silva, Walter; Marzocca, Piergiovanni; Librescu, Liviu

2001-01-01

An unified approach for dealing with stability and aeroelastic response to time-dependent pressure pulses of swept wings in an incompressible flow is developed. To this end the indicial function concept in time and frequency domains, enabling one to derive the proper unsteady aerodynamic loads is used. Results regarding stability in the frequency and time domains, and subcritical aeroelastic response to arbitrary time-dependent external excitation obtained via the direct use of the unsteady aerodynamic derivatives for 3-D wings are supplied. Closed form expressions for unsteady aerodynamic derivatives using this unified approach have been derived and used to illustrate their application to flutter and aeroelastic response to blast and sonic-boom signatures. In this context, an original representation of the aeroelastic response in the phase space was presented and pertinent conclusions on the implications of some basic parameters have been outlined.

Demonstration of relaxed static stability on a commercial transport

NASA Technical Reports Server (NTRS)

Rising, J. J.; Davis, W. J.; Willey, C. S.; Cokeley, R. C.

1984-01-01

Increasing jet aircraft fuel costs from 25 percent to nearly 60 percent of the aircraft direct operating costs have led to a heavy emphasis on the development of transport aircraft with significantly improved aerodynamic performance. The application of the concept of relaxed static stability (RSS) and the utilization of an active control stability augmentation system make it possible to design an aircraft with reduced aerodynamic trim drag due to a farther-aft cg balance. Reduced aerodynamic parasite drag and lower structural weight due to a smaller horizontal tail surface can also be obtained. The application of RSS has been studied under a NASA-sponsored program to determine ways of improving the energy efficiency in current and future transport aircraft. Attention is given to a near-term pitch active control system, an advanced pitch active control system, and an operational overview.

A General Theory of Unsteady Compressible Potential Aerodynamics

NASA Technical Reports Server (NTRS)

Morino, L.

1974-01-01

The general theory of potential aerodynamic flow around a lifting body having arbitrary shape and motion is presented. By using the Green function method, an integral representation for the potential is obtained for both supersonic and subsonic flow. Under small perturbation assumption, the potential at any point, P, in the field depends only upon the values of the potential and its normal derivative on the surface, sigma, of the body. Hence, if the point P approaches the surface of the body, the representation reduces to an integro-differential equation relating the potential and its normal derivative (which is known from the boundary conditions) on the surface sigma. For the important practical case of small harmonic oscillation around a rest position, the equation reduces to a two-dimensional Fredholm integral equation of second-type. It is shown that this equation reduces properly to the lifting surface theories as well as other classical mathematical formulas. The question of uniqueness is examined and it is shown that, for thin wings, the operator becomes singular as the thickness approaches zero. This fact may yield numerical problems for very thin wings.

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.

Nonplanar Method for Predicting Incompressible Aerodynamic Coefficients of Rectangular Wings with Circular-Arc Camber. Ph.D. Thesis - Virginia Polytechnic Institute

NASA Technical Reports Server (NTRS)

Lamar, J. E.

1971-01-01

The development of a nonplanar lifting surface method having a continuous distribution of singularities and satisfying the tangent flow boundary condition on the mean camber surface is given. The method predicts some incompressible longitudinal aerodynamic coefficients of rectangular wings which have circular-arc camber. The solution method is of the integral-equation type and the resulting surface integrals are evaluated by either using numerical or analytical techniques, as are appropriate. Applications are made and the results compared with those from an exact two-dimensional circular-arc camber solution, a three-dimensional flat-wing solution which represents the camber by a projected slope onto the flat surface, and a flat-wing experiment. From these comparisons, the present method is found to predict well the flat-wing experiment and limiting values, in addition to the center of pressure variation at an angle of attack of zero for any camber. For wings having camber ratios larger than about 1.25% and moderate to high aspect ratios, the results deterioriate due to the inadequacy of lifting pressure modes employed.

Estimated Benefits of Variable-Geometry Wing Camber Control for Transport Aircraft

NASA Technical Reports Server (NTRS)

Bolonkin, Alexander; Gilyard, Glenn B.

1999-01-01

Analytical benefits of variable-camber capability on subsonic transport aircraft are explored. Using aerodynamic performance models, including drag as a function of deflection angle for control surfaces of interest, optimal performance benefits of variable camber are calculated. Results demonstrate that if all wing trailing-edge surfaces are available for optimization, drag can be significantly reduced at most points within the flight envelope. The optimization approach developed and illustrated for flight uses variable camber for optimization of aerodynamic efficiency (maximizing the lift-to-drag ratio). Most transport aircraft have significant latent capability in this area. Wing camber control that can affect performance optimization for transport aircraft includes symmetric use of ailerons and flaps. In this paper, drag characteristics for aileron and flap deflections are computed based on analytical and wind-tunnel data. All calculations based on predictions for the subject aircraft and the optimal surface deflection are obtained by simple interpolation for given conditions. An algorithm is also presented for computation of optimal surface deflection for given conditions. Benefits of variable camber for a transport configuration using a simple trailing-edge control surface system can approach more than 10 percent, especially for nonstandard flight conditions. In the cruise regime, the benefit is 1-3 percent.

Feather roughness reduces flow separation during low Reynolds number glides of swifts.

PubMed

van Bokhorst, Evelien; de Kat, Roeland; Elsinga, Gerrit E; Lentink, David

2015-10-01

Swifts are aerodynamically sophisticated birds with a small arm and large hand wing that provides them with exquisite control over their glide performance. However, their hand wings have a seemingly unsophisticated surface roughness that is poised to disturb flow. This roughness of about 2% chord length is formed by the valleys and ridges of overlapping primary feathers with thick protruding rachides, which make the wing stiffer. An earlier flow study of laminar-turbulent boundary layer transition over prepared swift wings suggested that swifts can attain laminar flow at a low angle of attack. In contrast, aerodynamic design theory suggests that airfoils must be extremely smooth to attain such laminar flow. In hummingbirds, which have similarly rough wings, flow measurements on a 3D printed model suggest that the flow separates at the leading edge and becomes turbulent well above the rachis bumps in a detached shear layer. The aerodynamic function of wing roughness in small birds is, therefore, not fully understood. Here, we performed particle image velocimetry and force measurements to compare smooth versus rough 3D-printed models of the swift hand wing. The high-resolution boundary layer measurements show that the flow over rough wings is indeed laminar at a low angle of attack and a low Reynolds number, but becomes turbulent at higher values. In contrast, the boundary layer over the smooth wing forms open laminar separation bubbles that extend beyond the trailing edge. The boundary layer dynamics of the smooth surface varies non-linearly as a function of angle of attack and Reynolds number, whereas the rough surface boasts more consistent turbulent boundary layer dynamics. Comparison of the corresponding drag values, lift values and glide ratios suggests, however, that glide performance is equivalent. The increased structural performance, boundary layer robustness and equivalent aerodynamic performance of rough wings might have provided small (proto) birds with an evolutionary window to high glide performance. © 2015. Published by The Company of Biologists Ltd.

Simulation of upwind maneuvering of a sailing yacht

NASA Astrophysics Data System (ADS)

Harris, Daniel Hartrick

A time domain maneuvering simulation of an IACC class yacht suitable for the analysis of unsteady upwind sailing including tacking is presented. The simulation considers motions in six degrees of freedom. The hydrodynamic and aerodynamic loads are calculated primarily with unsteady potential theory supplemented by empirical viscous models. The hydrodynamic model includes the effects of incident waves. Control of the rudder is provided by a simple rate feedback autopilot which is augmented with open loop additions to mimic human steering. The hydrodynamic models are based on the superposition of force components. These components fall into two groups, those which the yacht will experience in calm water, and those due to incident waves. The calm water loads are further divided into zero Froude number, or "double body" maneuvering loads, hydrostatic loads, gravitational loads, free surface radiation loads, and viscous/residual loads. The maneuvering loads are calculated with an unsteady panel code which treats the instantaneous geometry of the yacht below the undisturbed free surface. The free surface radiation loads are calculated via convolution of impulse response functions derived from seakeeping strip theory. The viscous/residual loads are based upon empirical estimates. The aerodynamic model consists primarily of a database of steady state sail coefficients. These coefficients treat the individual contributions to the total sail force of a number of chordwise strips on both the main and jib. Dynamic effects are modeled by using the instantaneous incident wind velocity and direction as the independent variables for the sail load contribution of each strip. The sail coefficient database was calculated numerically with potential methods and simple empirical viscous corrections. Additional aerodynamic load calculations are made to determine the parasitic contributions of the rig and hull. Validation studies compare the steady sailing hydro and aerodynamic loads, seaway induced motions, added resistance in waves, and tacking performance with trials data and other sources. Reasonable agreement is found in all cases.

Protection of surface assets on Mars from wind blown jettisoned spacecraft components

NASA Astrophysics Data System (ADS)

Paton, Mark

2017-07-01

Jettisoned Entry, Descent and Landing System (EDLS) hardware from landing spacecraft have been observed by orbiting spacecraft, strewn over the Martian surface. Future Mars missions that land spacecraft close to prelanded assets will have to use a landing architecture that somehow minimises the possibility of impacts from these jettisoned EDLS components. Computer modelling is used here to investigate the influence of wind speed and direction on the distribution of EDLS components on the surface. Typical wind speeds encountered in the Martian Planetary Boundary Layer (PBL) were found to be of sufficient strength to blow items having a low ballistic coefficient, i.e. Hypersonic Inflatable Aerodynamic Decelerators (HIADs) or parachutes, onto prelanded assets even when the lander itself touches down several kilometres away. Employing meteorological measurements and careful characterisation of the Martian PBL, e.g. appropriate wind speed probability density functions, may then benefit future spacecraft landings, increase safety and possibly help reduce the delta v budget for Mars landers that rely on aerodynamic decelerators.

Wind tunnel test of a tailless aircraft with a belly-flap control surface used in combination with wing flaps

NASA Astrophysics Data System (ADS)

Dougherty, Daniel A.

A wind tunnel tat of a tailless aircraft configuration that has been quipped with a belly-flap control surface, was conducted with the goal of improving the trimmed maximum-lift coefficient. Tailless aircraft have aerodynamic and structural efficiencies that are superior to those of a traditionally configured wing/body/tail aircraft. However, tailless aircraft have a low maximum-lift coefficient such that; when sized for equivalent takeoff performance, the tailless aircraft suffers a large reduction in aerodynamic and structural efficiencies. A Belly-Flap control surface used in combination with wing trailing edge flaps was tested in a wind tunnel with the goal of achieving a longitudinally trimmed solution at a higher maximum lift coefficient. It was determined that, though the Belly-Flap increases the trimmed lift of the tailless configuration at low angles of attack, the maximum lift coefficient is slightly reduced in relation to the controls neutral configuration.

Optimization of the Upper Surface of Hypersonic Vehicle Based on CFD Analysis

NASA Astrophysics Data System (ADS)

Gao, T. Y.; Cui, K.; Hu, S. C.; Wang, X. P.; Yang, G. W.

2011-09-01

For the hypersonic vehicle, the aerodynamic performance becomes more intensive. Therefore, it is a significant event to optimize the shape of the hypersonic vehicle to achieve the project demands. It is a key technology to promote the performance of the hypersonic vehicle with the method of shape optimization. Based on the existing vehicle, the optimization to the upper surface of the Simplified hypersonic vehicle was done to obtain a shape which suits the project demand. At the cruising condition, the upper surface was parameterized with the B-Spline curve method. The incremental parametric method and the reconstruction technology of the local mesh were applied here. The whole flow field was been calculated and the aerodynamic performance of the craft were obtained by the computational fluid dynamic (CFD) technology. Then the vehicle shape was optimized to achieve the maximum lift-drag ratio at attack angle 3°, 4° and 5°. The results will provide the reference for the practical design.

Comparison of methods for the determination of NO-O3-NO2 fluxes and chemical interactions over a bare soil

NASA Astrophysics Data System (ADS)

Stella, P.; Loubet, B.; Laville, P.; Lamaud, E.; Cazaunau, M.; Laufs, S.; Bernard, F.; Grosselin, B.; Mascher, N.; Kurtenbach, R.; Mellouki, A.; Kleffmann, J.; Cellier, P.

2012-06-01

Tropospheric ozone (O3) is a known greenhouse gas responsible for impacts on human and animal health and ecosystem functioning. In addition, O3 plays an important role in tropospheric chemistry, together with nitrogen oxides. The determination of surface-atmosphere exchange fluxes of these trace gases is a prerequisite to establish their atmospheric budget and evaluate their impact onto the biosphere. In this study, O3, nitric oxide (NO) and nitrogen dioxide (NO2) fluxes were measured using the aerodynamic gradient method over a bare soil in an agricultural field. Ozone and NO fluxes were also measured using eddy-covariance and automatic chambers, respectively. The aerodynamic gradient measurement system, composed of fast response sensors, was capable to measure significant differences in NO and O3 mixing ratios between heights. However, due to local advection, NO2 mixing ratios were highly non-stationary and NO2 fluxes were, therefore, not significantly different from zero. The chemical reactions between O3, NO and NO2 led to little ozone flux divergence between the surface and the measurement height (less than 1% of the flux on average), whereas the NO flux divergence was about 10% on average. The use of fast response sensors allowed reducing the flux uncertainty. The aerodynamic gradient and the eddy-covariance methods gave comparable O3 fluxes. The chamber NO fluxes were down to 70% lower than the aerodynamic gradient fluxes, probably because of either the spatial heterogeneity of the soil NO emissions or the perturbation due to the chamber itself.

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.

Variable stiffness corrugated composite structure with shape memory polymer for morphing skin applications

NASA Astrophysics Data System (ADS)

Gong, Xiaobo; Liu, Liwu; Scarpa, Fabrizio; Leng, Jinsong; Liu, Yanju

2017-03-01

This work presents a variable stiffness corrugated structure based on a shape memory polymer (SMP) composite with corrugated laminates as reinforcement that shows smooth aerodynamic surface, extreme mechanical anisotropy and variable stiffness for potential morphing skin applications. The smart composite corrugated structure shows a low in-plane stiffness to minimize the actuation energy, but also possess high out-of-plane stiffness to transfer the aerodynamic pressure load. The skin provides an external smooth aerodynamic surface because of the one-sided filling with the SMP. Due to variable stiffness of the shape memory polymer the morphing skin exhibits a variable stiffness with a change of temperature, which can help the skin adjust its stiffness according different service environments and also lock the temporary shape without external force. Analytical models related to the transverse and bending stiffness are derived and validated using finite element techniques. The stiffness of the morphing skin is further investigated by performing a parametric analysis against the geometry of the corrugation and various sets of SMP fillers. The theoretical and numerical models show a good agreement and demonstrate the potential of this morphing skin concept for morphing aircraft applications. We also perform a feasibility study of the use of this morphing skin in a variable camber morphing wing baseline. The results show that the morphing skin concept exhibits sufficient bending stiffness to withstand the aerodynamic load at low speed (less than 0.3 Ma), while demonstrating a large transverse stiffness variation (up to 191 times) that helps to create a maximum mechanical efficiency of the structure under varying external conditions.

Aerodynamic characteristics of flying fish in gliding flight.

PubMed

Park, Hyungmin; Choi, Haecheon

2010-10-01

The flying fish (family Exocoetidae) is an exceptional marine flying vertebrate, utilizing the advantages of moving in two different media, i.e. swimming in water and flying in air. Despite some physical limitations by moving in both water and air, the flying fish has evolved to have good aerodynamic designs (such as the hypertrophied fins and cylindrical body with a ventrally flattened surface) for proficient gliding flight. Hence, the morphological and behavioral adaptations of flying fish to aerial locomotion have attracted great interest from various fields including biology and aerodynamics. Several aspects of the flight of flying fish have been determined or conjectured from previous field observations and measurements of morphometric parameters. However, the detailed measurement of wing performance associated with its morphometry for identifying the characteristics of flight in flying fish has not been performed yet. Therefore, in the present study, we directly measure the aerodynamic forces and moment on darkedged-wing flying fish (Cypselurus hiraii) models and correlated them with morphological characteristics of wing (fin). The model configurations considered are: (1) both the pectoral and pelvic fins spread out, (2) only the pectoral fins spread with the pelvic fins folded, and (3) both fins folded. The role of the pelvic fins was found to increase the lift force and lift-to-drag ratio, which is confirmed by the jet-like flow structure existing between the pectoral and pelvic fins. With both the pectoral and pelvic fins spread, the longitudinal static stability is also more enhanced than that with the pelvic fins folded. For cases 1 and 2, the lift-to-drag ratio was maximum at attack angles of around 0 deg, where the attack angle is the angle between the longitudinal body axis and the flying direction. The lift coefficient is largest at attack angles around 30∼35 deg, at which the flying fish is observed to emerge from the sea surface. From glide polar, we find that the gliding performance of flying fish is comparable to those of bird wings such as the hawk, petrel and wood duck. However, the induced drag by strong wing-tip vortices is one of the dominant drag components. Finally, we examine ground effect on the aerodynamic forces of the gliding flying fish and find that the flying fish achieves the reduction of drag and increase of lift-to-drag ratio by flying close to the sea surface.

Aerodynamic coefficients in generalized unsteady thin airfoil theory

NASA Technical Reports Server (NTRS)

Williams, M. H.

1980-01-01

Two cases are considered: (1) rigid body motion of an airfoil-flap combination consisting of vertical translation of given amplitude, rotation of given amplitude about a specified axis, and rotation of given amplitude of the control surface alone about its hinge; the upwash for this problem is defined mathematically; and (2) sinusoidal gust of given amplitude and wave number, for which the upwash is defined mathematically. Simple universal formulas are presented for the most important aerodynamic coefficients in unsteady thin airfoil theory. The lift and moment induced by a generalized gust are evaluated explicitly in terms of the gust wavelength. Similarly, in the control surface problem, the lift, moment, and hinge moments are given as explicit algebraic functions of hinge location. These results can be used together with any of the standard numerical inversion routines for the elementary loads (pitch and heave).

Biomimetic optimization research on wind noise reduction of an asymmetric cross-section bar.

PubMed

Zhang, Yingchao; Meng, Weijiang; Fan, Bing; Tang, Wenhui

2016-01-01

In this paper, we used the principle of biomimetics to design two-dimensional and three-dimensional bar sections, and used computational fluid dynamics software to numerically simulate and analyse the aerodynamic noise, to reduce drag and noise. We used the principle of biomimetics to design the cross-section of a bar. An owl wing shape was used for the initial design of the section geometry; then the feathered form of an owl wing, the v-shaped micro-grooves of a shark's skin, the tubercles of a humpback whale's flipper, and the stripy surface of a scallop's shell were used to inspire surface features, added to the initial section and three-dimensional shape. Through computational aeroacoustic simulations, we obtained the aerodynamic characteristics and the noise levels of the models. These biomimetic models dramatically decreased noise levels.

A Rapid Aerodynamic Design Procedure Based on Artificial Neural Networks

NASA Technical Reports Server (NTRS)

Rai, Man Mohan

2001-01-01

An aerodynamic design procedure that uses neural networks to model the functional behavior of the objective function in design space has been developed. This method incorporates several improvements to an earlier method that employed a strategy called parameter-based partitioning of the design space in order to reduce the computational costs associated with design optimization. As with the earlier method, the current method uses a sequence of response surfaces to traverse the design space in search of the optimal solution. The new method yields significant reductions in computational costs by using composite response surfaces with better generalization capabilities and by exploiting synergies between the optimization method and the simulation codes used to generate the training data. These reductions in design optimization costs are demonstrated for a turbine airfoil design study where a generic shape is evolved into an optimal airfoil.

The effects of inlet turbulence and rotor/stator interactions on the aerodynamics and heat transfer of a large-scale rotating turbine model, volume 1

NASA Technical Reports Server (NTRS)

Dring, R. P.; Blair, M. F.; Joslyn, H. D.; Power, G. D.; Verdon, J. M.

1987-01-01

A combined experimental and analytical program was conducted to examine the effects of inlet turbulence on airfoil heat transfer. Heat transfer measurements were obtained using low conductivity airfoils with miniature thermocouples welded to a thin, electrically heated surface skin. Heat transfer data were acquired for various combinations of low or high inlet turbulence intensity, flow coefficient (incidence), first-stator/rotor axial spacing, Reynolds number, and relative circumferential position of the first and second stators. Aerodynamic measurements include distributions of the mean and fluctuating velocities at the turbine inlet and, for each airfoil row, midspan airfoil surface pressures and circumferential distributions of the downstream steady state pressures and fluctuating velocities. Analytical results include airfoil heat transfer predictions and a examination of solutions of the unstead boundary layer equipment.

Unconventional missile concepts from consideration of varied mission requirements

NASA Technical Reports Server (NTRS)

Spearman, M. L.

1984-01-01

Missile concepts for volumetric efficiency, minimum carriage constraints, and aerodynamic performance to achieve mission requirements. The mission requirements considered include air to surface roles such as defense suppression or antishipping where payload and range may have priority over high maneuver capability, and air to air and surface to air roles paying attention to good maneuvering capability. The concepts are intended to provide for ease of storage or carriage. The concepts include monoplanes with highly swept, thick delta wings, highly swept delta wings mounted either high or low on a semicircular body, some ring wing and semiring wing arrangements, parasol wing, and elliptical lifting bodies. The missile configurations indicate possible approaches toward resolving problems of carriage and storage while retaining good volumetric and aerodynamic efficiency. The configurations can accomplish a variety of possible missions with relatively simple vehicle shapes.

Aerodynamic performance of high turning core turbine vanes in a two dimensional cascade

NASA Technical Reports Server (NTRS)

Schwab, J. R.

1982-01-01

Experimental and theoretical aerodynamic performance data are presented for four uncooled high turning core turbine vanes with exit angles of 74.9, 75.0, 77.5, and 79.6 degrees in a two dimensional cascade. Data for a more conservative 67.0 degree vane are included for comparison. Correction of the experimental aftermix kinetic energy losses to a common 0.100 centimeter trailing edge thickness yields a linear trend of increased loss from 0.020 to 0.025 as the vane exit angle increases from 67.0 to 79.6 degrees. The theoretical losses show a similar trend. The experimental and theoretical vane surface velocity distributions generally agree within approximately five percent, although the suction surface theoretical velocities are generally higher than the experimental velocities as the vane exit angle increases.

Aerodynamic heating rate distributions induced by trailing edge controls on hypersonic aircraft configurations at Mach 8

NASA Technical Reports Server (NTRS)

Kaufman, L. G., II; Johnson, C. B.

1984-01-01

Aerodynamic surface heating rate distributions in three dimensional shock wave boundary layer interaction flow regions are presented for a generic set of model configurations representative of the aft portion of hypersonic aircraft. Heat transfer data were obtained using the phase change coating technique (paint) and, at particular spanwise and streamwise stations for sample cases, by the thin wall transient temperature technique (thermocouples). Surface oil flow patterns are also shown. The good accuracy of the detailed heat transfer data, as attested in part by their repeatability, is attributable partially to the comparatively high temperature potential of the NASA-Langley Mach 8 Variable Density Tunnel. The data are well suited to help guide heating analyses of Mach 8 aircraft, and should be considered in formulating improvements to empiric analytic methods for calculating heat transfer rate coefficient distributions.

A study of aerodynamic heating distributions on a tip-fin controller installed on a Space Shuttle Orbiter model

NASA Technical Reports Server (NTRS)

Wittliff, C. E.

1982-01-01

The aerodynamic heating of a tip-fin controller mounted on a Space Shuttle Orbiter model was studied experimentally in the Calspan Advanced Technology Center 96 inch Hypersonic Shock Tunnel. A 0.0175 scale model was tested at Mach numbers from 10 to 17.5 at angles of attack typical of a shuttle entry. The study was conducted in two phases. In phase 1 testing a thermographic phosphor technique was used to qualitatively determine the areas of high heat-transfer rates. Based on the results of this phase, the model was instrumented with 40 thin-film resistance thermometers to obtain quantitative measurements of the aerodynamic heating. The results of the phase 2 testing indicate that the highest heating rates, which occur on the leading edge of the tip-fin controller, are very sensitive to angle of attack for alpha or = 30 deg. The shock wave from the leading edge of the orbiter wing impinges on the leading edge of the tip-fin controller resulting in peak values of h/h(Ref) in the range from 1.5 to 2.0. Away from the leading edge, the heat-transfer rates never exceed h/h(Ref) = 0.25 when the control surface, is not deflected. With the control surface deflected 20 deg, the heat-transfer rates had a maximum value of h/h(Ref) = 0.3. The heating rates are quite nonuniform over the outboard surface and are sensitive to angle of attack.

Application of the aerodynamic energy concept to flutter suppression and gust alleviation by use of active controls

NASA Technical Reports Server (NTRS)

Nissim, E.; Caspi, A.; Lottati, I.

1976-01-01

The effects of active controls on flutter suppression and gust alleviation of the Arava twin turboprop STOL transport and the Westwind twinjet business transport are investigated. The active control surfaces are introduced in pairs which include, in any chosen wing strip, a 20-percent chord leading-edge control and a 20-percent chord trailing-edge control. Each control surface is driven by a combined linear-rotational sensor system located on the activated strip. The control law is based on the concept of aerodynamic energy and utilizes previously optimized control law parameters based on two-dimensional aerodynamic theory. The best locations of the activated system along the span of the wing are determined for bending-moment alleviation, reduction in fuselage accelerations, and flutter suppression. The effectiveness of the activated system over a wide range of maximum control deflections is also determined. Two control laws are investigated. The first control law utilizes both rigid-body and elastic contributions of the motion. The second control law employs primarily the elastic contribution of the wing and leads to large increases in the activated control effectiveness as compared with the basic control law. The results indicate that flutter speed can be significantly increased (over 70 percent increase) and that the bending moment due to gust loading can be almost totally eliminated by a control system of about 10 to 20 percent span with reasonable control-surface rotations.

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.

Does surface roughness dominate biophysical forcing of land use and land cover change in the eastern United States?

NASA Astrophysics Data System (ADS)

Burakowski, E. A.; Tawfik, A. B.; Ouimette, A.; Lepine, L. C.; Ollinger, S. V.; Bonan, G. B.; Zarzycki, C. M.; Novick, K. A.

2016-12-01

Changes in land use, land cover, or both promote changes in surface temperature that can amplify or dampen long-term trends driven by natural and anthropogenic climate change by modifying the surface energy budget, primarily through differences in albedo, evapotranspiration, and aerodynamic roughness. Recent advances in variable resolution global models provide the tools necessary to investigate local and global impacts of land use and land cover change by embedding a high-resolution grid over areas of interest in a seamless and computationally efficient manner. Here, we used two eddy covariance tower clusters in the Eastern US (University of New Hampshire UNH and Duke Forest) to validate simulation of surface energy fluxes and properties by the uncoupled Community Land Model (PTCLM4.5) and coupled land-atmosphere Variable-Resolution Community Earth System Model (VR-CESM1.3). Surface energy fluxes and properties are generally well captured by the models for grassland sites, however forested sites tend to underestimate latent heat and overestimate sensible heat flux. Surface roughness emerged as the dominant biophysical forcing factor affecting surface temperature in the eastern United States, generally leading to warmer nighttime temperatures and cooler daytime temperatures. However, the sign and magnitude of the roughness effect on surface temperature was highly sensitive to the calculation of aerodynamic resistance to heat transfer.

Wind-Tunnel Tests of a Portion of a PV-2 Helicopter Rotor Blade

NASA Technical Reports Server (NTRS)

Kemp, William B., Jr.

1945-01-01

A portion of a PV-2 helicopter rotor blade has been tested in the 6- by 6-foot test section of the Langley stability tunnel to determine if the aerodynamic characteristics were seriously affected by cross flow or fabric distortion. The outer portion of the blade was tested as a reflection plane model pivoted about the tunnel wall to obtain various angles of cross flow over the blade. Because the tunnel wall acts as a plane of sytry, the measured aerodynamic characteristics correspond to those of an airfoil having various angles of sweepforward and sweepback. Tests were made with the vents on the lower surface open and also with the vents sealed and the internal pressure held at -20 inches of water producing an internal pressure coefficient of -1.059. The change in contour resulting from the range of internal pressures used had very little effect on the aerodynamic characteristics of the blade. The test methods were considered to simulate inadequately the flow conditions over the rotor blade because the effects of cross flow were limited to conditions corresponding to sweep of the blade. The results indicated that this type of cross flow had only minor effects on the aerodynamic characteristics of the blade. It is believed, therefore, that future tests to determine the effects on the aerodynamic characteristics of cross flow should utilize complete rotors.

Mars 2020 Entry, Descent, and Landing Instrumentation 2 (MEDLI2) Sensor Suite

NASA Technical Reports Server (NTRS)

Hwang, Helen; Wright, Henry; Kuhl, Chris; Schoenenberger, Mark; White, Todd; Karlgaard, Chris; Mahzari, Milad; Oishi, Tomo; Pennington, Steve; Trombetta, Nick;

Results of a space shuttle vehicle ferry configuration afterbody fairing optimization study using a 140A/B 0.0405-scale model orbiter (43-0) in the Rockwell International 7.75 by 11.0 ft low speed wind tunnel (OA124)

NASA Technical Reports Server (NTRS)

Houlihan, S. R.

1975-01-01

Experimental aerodynamic investigations were conducted on a dual-strut mounted 0.0405-scale representation of the 140A/B outer mold line space shuttle orbiter vehicle. The tests, conducted from 11 Oct., 1974 through 22 Oct., 1974, were primarily to investigate aerodynamic stability and control characteristics of the space shuttle orbiter ferry configuration. Four afterbody fairing configurations and various additions to them in the form of horizontal and ventral fins strakes and other aerodynamic protuberances were tested. Base line data on the basic orbiter with MPS nozzles and bodyflap were recorded. The drag of the optimum ferry configuration was increased to the level of the basic orbiter for possible flight test configurations by the addition of two sizes of perforated speed brakes on the tail cone surface.

Results of tests OA12 and IA9 in the Ames Research Center unitary plan wind tunnels on an 0.030-scale model of the space shuttle vehicle 2A to determine aerodynamic loads, volume 1

NASA Technical Reports Server (NTRS)

Spangler, R. H.

1973-01-01

Tests were conducted in unitary plan wind tunnels on an 0.030-scale replica of the space shuttle vehicle configuration 2A. Aerodynamic loads data were obtained at Mach numbers from 0.6 to 3.5. The investigation included tests on the integrated (launch) configuration and tests on the isolated orbiter (entry configuration). The integrated vehicle was tested at angles of attack and sideslip from minus 8 deg to plus 8 deg. The isolated orbiter was tested at angles of attack from minus 15 deg to plus 40 deg and angles of sideslip from minus 10 deg to plus 10 deg are dictated by trajectory considerations. The effects of orbiter/external tank incidence and deflected control surfaces on aerodynamic loads were also investigated.

Results of tests OA12 and IA9 in the Ames Research Center unitary plan wind tunnels on an 0.030 scale model of the space shuttle vehicle 2A to determine aerodynamic loads, volume 7

NASA Technical Reports Server (NTRS)

Spangler, R. H.

1973-01-01

Tests were conducted in wind tunnels during April and May 1973, on an 0.030-scale replica of the Space Shuttle Vehicle Configuration 2A. Aerodynamic loads data were obtained at Mach numbers from 0.6 to 3.5. The investigation included tests on the integrated (launch) configuration and on the isolated orbiter (entry configuration). The integrated vehicle was tested at angles of attack and sideslip from -8 deg. The isolated orbiter was tested at angles of attack from -15 deg to +40 deg and angles of sideslip from -10 deg to +10 deg as dictated by trajectory considerations. The effects of orbiter/external tank incidence angle and deflected control surfaces on aerodynamic loads were also investigated.

Wing Weight Optimization Under Aeroelastic Loads Subject to Stress Constraints

NASA Technical Reports Server (NTRS)

Kapania, Rakesh K.; Issac, J.; Macmurdy, D.; Guruswamy, Guru P.

1997-01-01

A minimum weight optimization of the wing under aeroelastic loads subject to stress constraints is carried out. The loads for the optimization are based on aeroelastic trim. The design variables are the thickness of the wing skins and planform variables. The composite plate structural model incorporates first-order shear deformation theory, the wing deflections are expressed using Chebyshev polynomials and a Rayleigh-Ritz procedure is adopted for the structural formulation. The aerodynamic pressures provided by the aerodynamic code at a discrete number of grid points is represented as a bilinear distribution on the composite plate code to solve for the deflections and stresses in the wing. The lifting-surface aerodynamic code FAST is presently being used to generate the pressure distribution over the wing. The envisioned ENSAERO/Plate is an aeroelastic analysis code which combines ENSAERO version 3.0 (for analysis of wing-body configurations) with the composite plate code.

Rotary-wing aerodynamics. Volume 1: Basic theories of rotor aerodynamics with application to helicopters. [momentum, vortices, and potential theory

NASA Technical Reports Server (NTRS)

Stepniewski, W. Z.

1979-01-01

The concept of rotary-wing aircraft in general is defined. The energy effectiveness of helicopters is compared with that of other static thrust generators in hover, as well as with various air and ground vehicles in forward translation. The most important aspects of rotor-blade dynamics and rotor control are reviewed. The simple physicomathematical model of the rotor offered by the momentum theory is introduced and its usefulness and limitations are assessed. The combined blade-element and momentum theory approach, which provides greater accuracy in performance predictions, is described as well as the vortex theory which models a rotor blade by means of a vortex filament or vorticity surface. The application of the velocity and acceleration potential theory to the determination of flow fields around three dimensional, non-rotating bodies as well as to rotor aerodynamic problems is described. Airfoil sections suitable for rotors are also considered.

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.

Constraints on the wing morphology of pterosaurs

PubMed Central

Palmer, Colin; Dyke, Gareth

2012-01-01

Animals that fly must be able to do so over a huge range of aerodynamic conditions, determined by weather, wind speed and the nature of their environment. No single parameter can be used to determine—let alone measure—optimum flight performance as it relates to wing shape. Reconstructing the wings of the extinct pterosaurs has therefore proved especially problematic: these Mesozoic flying reptiles had a soft-tissue membranous flight surface that is rarely preserved in the fossil record. Here, we review basic mechanical and aerodynamic constraints that influenced the wing shape of pterosaurs, and, building on this, present a series of theoretical modelling results. These results allow us to predict the most likely wing shapes that could have been employed by these ancient reptiles, and further show that a combination of anterior sweep and a reflexed proximal wing section provides an aerodynamically balanced and efficient theoretical pterosaur wing shape, with clear benefits for their flight stability. PMID:21957137

Enhancement of the CAVE computer code. [aerodynamic heating package for nose cones and scramjet engine sidewalls

NASA Technical Reports Server (NTRS)

Rathjen, K. A.; Burk, H. O.

1983-01-01

The computer code CAVE (Conduction Analysis via Eigenvalues) is a convenient and efficient computer code for predicting two dimensional temperature histories within thermal protection systems for hypersonic vehicles. The capabilities of CAVE were enhanced by incorporation of the following features into the code: real gas effects in the aerodynamic heating predictions, geometry and aerodynamic heating package for analyses of cone shaped bodies, input option to change from laminar to turbulent heating predictions on leading edges, modification to account for reduction in adiabatic wall temperature with increase in leading sweep, geometry package for two dimensional scramjet engine sidewall, with an option for heat transfer to external and internal surfaces, print out modification to provide tables of select temperatures for plotting and storage, and modifications to the radiation calculation procedure to eliminate temperature oscillations induced by high heating rates. These new features are described.

Highly variable aerodynamic roughness length (z0) for a hummocky debris-covered glacier

NASA Astrophysics Data System (ADS)

Miles, Evan S.; Steiner, Jakob F.; Brun, Fanny

2017-08-01

The aerodynamic roughness length (z0) is an essential parameter in surface energy balance studies, but few literature values exist for debris-covered glaciers. We use microtopographic and aerodynamic methods to assess the spatial variability of z0 for Lirung Glacier, Nepal. We apply structure from motion to produce digital elevation models for three nested domains: five 1 m2 plots, a 21,300 m2 surface depression, and the lower 550,000 m2 of the debris-mantled tongue. Wind and temperature sensor towers were installed in the vicinity of the plots within the surface depression in October 2014. We calculate z0 according to a variety of transect-based microtopographic parameterizations for each plot, then develop a grid version of the algorithms by aggregating data from all transects. This grid approach is applied to the surface depression digital elevation model to characterize z0 spatial variability. The algorithms reproduce the same variability among transects and plots, but z0 estimates vary by an order of magnitude between algorithms. Across the study depression, results from different algorithms are strongly correlated. Using Monin-Obukov similarity theory, we derive z0 values from the meteorological data. Using different stability criteria, we derive median values of z0 between 0.03 m and 0.05 m, but with considerable uncertainty due to the glacier's complex topography. Considering estimates from these algorithms, results suggest that z0 varies across Lirung Glacier between ˜0.005 m (gravels) to ˜0.5 m (boulders). Future efforts should assess the importance of such variable z0 values in a distributed energy balance model.

Plasma Flowfields Around Low Earth Orbit Objects: Aerodynamics to Underpin Orbit Predictions

NASA Astrophysics Data System (ADS)

Capon, Christopher; Boyce, Russell; Brown, Melrose

2016-07-01

Interactions between orbiting bodies and the charged space environment are complex. The large variation in passive body parameters e.g. size, geometry and materials, makes the plasma-body interaction in Low Earth Orbit (LEO) a region rich in fundamental physical phenomena. The aerodynamic interaction of LEO orbiting bodies with the neutral environment constitutes the largest non-conservative force on the body. However in general, study of the LEO plasma-body interaction has not been concerned with external flow physics, but rather with the effects on surface charging. The impact of ionospheric flow physics on the forces on space debris (and active objects) is not well understood. The work presented here investigates the contribution that plasma-body interactions have on the flow structure and hence on the total atmospheric force vector experienced by a polar orbiting LEO body. This work applies a hybrid Particle-in-Cell (PIC) - Direct Simulation Monte Carlo (DSMC) code, pdFoam, to self-consistently model the electrostatic flowfield about a cylinder with a uniform, fixed surface potential. Flow conditions are representative of the mean conditions experienced by the Earth Observing Satellite (EOS) based on the International Reference Ionosphere model (IRI-86). The electron distribution function is represented by a non-linear Boltzmann electron fluid and ion gas-surface interactions are assumed to be that of a neutralising, conducting, thermally accommodating solid wall with diffuse reflections. The variation in flowfield and aerodynamic properties with surface potential at a fixed flow condition is investigated, and insight into the relative contributions of charged and neutral species to the flow physics experienced by a LEO orbiting body is provided. This in turn is intended to help improve the fidelity of physics-based orbit predictions for space debris and other near-Earth space objects.

Aerodynamic heating and the deflection of drops by an obstacle in an air stream in relation to aircraft icing

NASA Technical Reports Server (NTRS)

Kantrowitz, Arthur

1940-01-01

Two topics of interest to persons attempting to apply the heat method of preventing ice formation on aircraft are considered. Surfaces moving through air at high speed are shown, both theoretically and experimentally, to be subject to important aerodynamic heating effects that will materially reduce the heat required to prevent ice. Numerical calculations of the path of water drops in an air stream around a circular cylinder are given. From these calculations, information is obtained on the percentage of the swept area cleared of drops.

Actuator concepts and mechatronics

NASA Astrophysics Data System (ADS)

Gilbert, Michael G.; Horner, Garnett C.

1998-06-01

Mechatronic design implies the consideration of integrated mechanical, electrical, and local control characteristics in electromechanical device design. In this paper, mechatronic development of actuation device concepts for active aircraft aerodynamic flow control are presented and discussed. The devices are intended to be embedded in aircraft aerodynamic surfaces to provide zero-net-momentum jets or additional flow-vorticity to control boundary layers and flow- separation. Two synthetic jet device prototypes and one vorticity-on-demand prototype currently in development are described in the paper. The aspects of actuation materials, design approaches to generating jets and vorticity, and the integration of miniaturized electronics are stressed.

Numerical modeling of wind turbine aerodynamic noise in the time domain.

PubMed

Lee, Seunghoon; Lee, Seungmin; Lee, Soogab

2013-02-01

Aerodynamic noise from a wind turbine is numerically modeled in the time domain. An analytic trailing edge noise model is used to determine the unsteady pressure on the blade surface. The far-field noise due to the unsteady pressure is calculated using the acoustic analogy theory. By using a strip theory approach, the two-dimensional noise model is applied to rotating wind turbine blades. The numerical results indicate that, although the operating and atmospheric conditions are identical, the acoustical characteristics of wind turbine noise can be quite different with respect to the distance and direction from the wind turbine.

Study of aerodynamic surface control of space shuttle boost and reentry, volume 1

NASA Technical Reports Server (NTRS)

Chang, C. J.; Connor, C. L.; Gill, G. P.

1972-01-01

The optimization technique is described which was used in the study for applying modern optimal control technology to the design of shuttle booster engine reaction control systems and aerodynamic control systems. Complete formulations are presented for both the ascent and reentry portions of the study. These formulations include derivations of the 6D perturbation equations of motion and the process followed in the control and blending law selections. A total hybrid software concept applied to the study is described in detail. Conclusions and recommendations based on the results of the study are included.

Low-speed Aerodynamic Investigations of a Hybrid Wing Body Configuration

NASA Technical Reports Server (NTRS)

Vicroy, Dan D.; Gatlin, Gregory M.; Jenkins, Luther N.; Murphy, Patrick C.; Carter, Melissa B.

2014-01-01

Two low-speed static wind tunnel tests and a water tunnel static and dynamic forced-motion test have been conducted on a hybrid wing-body (HWB) twinjet configuration. These tests, in addition to computational fluid dynamics (CFD) analysis, have provided a comprehensive dataset of the low-speed aerodynamic characteristics of this nonproprietary configuration. In addition to force and moment measurements, the tests included surface pressures, flow visualization, and off-body particle image velocimetry measurements. This paper will summarize the results of these tests and highlight the data that is available for code comparison or additional analysis.

A one-layer satellite surface energy balance for estimating evapotranspiration rates and crop water stress indexes.

PubMed

Barbagallo, Salvatore; Consoli, Simona; Russo, Alfonso

2009-01-01

Daily evapotranspiration fluxes over the semi-arid Catania Plain area (Eastern Sicily, Italy) were evaluated using remotely sensed data from Landsat Thematic Mapper TM5 images. A one-source parameterization of the surface sensible heat flux exchange using satellite surface temperature has been used. The transfer of sensible and latent heat is described by aerodynamic resistance and surface resistance. Required model inputs are brightness, temperature, fractional vegetation cover or leaf area index, albedo, crop height, roughness lengths, net radiation, air temperature, air humidity and wind speed. The aerodynamic resistance (r(ah)) is formulated on the basis of the Monin-Obukhov surface layer similarity theory and the surface resistance (r(s)) is evaluated from the energy balance equation. The instantaneous surface flux values were converted into evaporative fraction (EF) over the heterogeneous land surface to derive daily evapotranspiration values. Remote sensing-based assessments of crop water stress (CWSI) were also made in order to identify local irrigation requirements. Evapotranspiration data and crop coefficient values obtained from the approach were compared with: (i) data from the semi-empirical approach "K(c) reflectance-based", which integrates satellite data in the visible and NIR regions of the electromagnetic spectrum with ground-based measurements and (ii) surface energy flux measurements collected from a micrometeorological tower located in the experiment area. The expected variability associated with ET flux measurements suggests that the approach-derived surface fluxes were in acceptable agreement with the observations.

Analysis and evaluation of an integrated laminar flow control propulsion system

NASA Technical Reports Server (NTRS)

Keith, Theo G., Jr.; Dewitt, Kenneth J.

1993-01-01

Reduction of drag has been a major goal of the aircraft industry as no other single quantity influences the operating costs of transport aircraft more than aerodynamic drag. It has been estimated that even modest reduction of frictional drag could reduce fuel costs by anywhere from 2 to 5 percent. Current research on boundary layer drag reduction deals with various approaches to reduce turbulent skin friction drag as a means of improving aircraft performance. One of the techniques belonging to this category is laminar flow control in which extensive regions of laminar flow are maintained over aircraft surfaces by delaying transition to turbulence through the ingestion of boundary layer air. While problems of laminar flow control have been studied in some detail, the prospect of improving the propulsion system of an aircraft by the use of ingested boundary layer air has received very little attention. An initial study for the purpose of reducing propulsion system requirements by utilizing the kinetic energy of boundary layer air was performed in the mid-1970's at LeRC. This study which was based on ingesting the boundary layer air at a single location, did not yield any significant overall propulsion benefits; therefore, the concept was not pursued further. However, since then it has been proposed that if the boundary layer air were ingested at various locations on the aircraft surface instead of just at one site, an improvement in the propulsion system might be realized. The present report provides a review of laminar flow control by suction and focuses on the problems of reducing skin friction drag by maintaining extensive regions of laminar flow over the aircraft surfaces. In addition, it includes an evaluation of an aircraft propulsion system that is augmented by ingested boundary layer air.

Redesigning of a Canard Control Surface of an Advanced Fighter Aircraft: Effect on Buckling and Aerodynamic Behavior

NASA Astrophysics Data System (ADS)

Shrivastava, Sachin; Mohite, P. M.

2015-01-01

A redesign of canard control-surface of an advanced all-metallic fighter aircraft was carried out by using carbon fibre composite (CFC) for ribs and panels. In this study ply-orientations of CFC structure are optimized using a Genetic-Algorithm (GA) with an objective function to have minimum failure index (FI) according to Tsai-Wu failure criterion. The redesigned CFC structure was sufficiently strong to withstand aerodynamic loads from stress and deflection points of view. Now, in the present work CFC canard structure has been studied for its buckling strength in comparison to existing metallic design. In this study, the existing metallic design was found to be weak in buckling. Upon a detailed investigation, it was revealed that there are reported failures in the vicinity of zones where initial buckling modes are excited as predicted by the finite element based buckling analysis. In view of buckling failures, the redesigned CFC structure is sufficiently reinforced with stringers at specific locations. After providing reinforcements against buckling, the twist and the camber variations of the airfoil are checked and compared with existing structure data. Finally, the modal analysis has been carried out to compare the variation in excitation frequency due to material change. The CFC structure thus redesigned is safe from buckling and aerodynamic aspects as well.

Research at NASA's NFAC wind tunnels

NASA Technical Reports Server (NTRS)

Edenborough, H. Kipling

1990-01-01

The National Full-Scale Aerodynamics Complex (NFAC) is a unique combination of wind tunnels that allow the testing of aerodynamic and dynamic models at full or large scale. It can even accommodate actual aircraft with their engines running. Maintaining full-scale Reynolds numbers and testing with surface irregularities, protuberances, and control surface gaps that either closely match the full-scale or indeed are those of the full-scale aircraft help produce test data that accurately predict what can be expected from future flight investigations. This complex has grown from the venerable 40- by 80-ft wind tunnel that has served for over 40 years helping researchers obtain data to better understand the aerodynamics of a wide range of aircraft from helicopters to the space shuttle. A recent modification to the tunnel expanded its maximum speed capabilities, added a new 80- by 120-ft test section and provided extensive acoustic treatment. The modification is certain to make the NFAC an even more useful facility for NASA's ongoing research activities. A brief background is presented on the original facility and the kind of testing that has been accomplished using it through the years. A summary of the modification project and the measured capabilities of the two test sections is followed by a review of recent testing activities and of research projected for the future.

Analysis of Ground Effects on Aerodynamic Characteristics of Aerofoils Using Boundary Layer Approximation

NASA Astrophysics Data System (ADS)

Takahashi, Yuji; Kikuchi, Masanori; Hirano, Kimitaka

A study of a new high-speed zero-emission transportation “Aerotrain” is being carried out in Tohoku University and the University of Miyazaki. Because the aerotrain utilizes the ground effect, research on the aerofoil section, which can harness the ground effect effectively, is important. The aerotrain moves along a U-shaped guideway, which has a ground and sidewalls, so it has many viscous interference elements. In an analysis of the ground effects on the aerodynamic characteristics of aerofoils, the boundary layers on the aerofoil surface must be considered. At first, velocity distributions on the surfaces of aerofoils in potential flows are computed using the vortex method, then the momentum integration equations of the boundary layer are solved with experimental formulas. This procedure has the following advantages: modifications of the aerofoil section are easy because it is not necessary to make complicated computational grids, boundary layer transition and separation can be predicted using empirical procedures. The aerodynamic characteristics of four types of aerofoil sections are investigated to clarify the relationship between aerofoil sections and ground effects. Computational results are compared with experimental results obtained using a towing wind tunnel to verify computational precisions. In addition, aerofoil characteristics at an actual cruise speed are analyzed.

Flight testing of live Monarch butterflies to determine the aerodynamic benefit of butterfly scales

NASA Astrophysics Data System (ADS)

Lang, Amy; Cranford, Jacob; Conway, Jasmine; Slegers, Nathan; Dechello, Nicole; Wilroy, Jacob

2014-11-01

Evolutionary adaptations in the morphological structure of butterfly scales (0.1 mm in size) to develop a unique micro-patterning resulting in a surface drag alteration, stem from a probable aerodynamic benefit of minimizing the energy requirement to fly a very lightweight body with comparably large surface area in a low Re flow regime. Live Monarch butterflies were tested at UAHuntsville's Autonomous Tracking and Optical Measurement (ATOM) Laboratory, which uses 22 Vicon T40 cameras that allow for millimeter level tracking of reflective markers at 515 fps over a 4 m × 6 m × 7 m volume. Data recorded included the flight path as well as the wing flapping angle and wing-beat frequency. Insects were first tested with their scales intact, and then again with the scales carefully removed. Differences in flapping frequency and/or energy obtained during flight due to the removal of the scales will be discussed. Initial data analysis indicates that scale removal in some specimens leads to increased flapping frequencies for similar energetic flight or reduced flight speed for similar flapping frequencies. Both results point to the scales providing an aerodynamic benefit, which is hypothesized to be linked to leading-edge vortex formation and induced drag. Funding from the National Science Foundation (CBET and REU) is gratefully acknowledged.

Precise and efficient evaluation of gravimetric quantities at arbitrarily scattered points in space

NASA Astrophysics Data System (ADS)

Ivanov, Kamen G.; Pavlis, Nikolaos K.; Petrushev, Pencho

2017-12-01

Gravimetric quantities are commonly represented in terms of high degree surface or solid spherical harmonics. After EGM2008, such expansions routinely extend to spherical harmonic degree 2190, which makes the computation of gravimetric quantities at a large number of arbitrarily scattered points in space using harmonic synthesis, a very computationally demanding process. We present here the development of an algorithm and its associated software for the efficient and precise evaluation of gravimetric quantities, represented in high degree solid spherical harmonics, at arbitrarily scattered points in the space exterior to the surface of the Earth. The new algorithm is based on representation of the quantities of interest in solid ellipsoidal harmonics and application of the tensor product trigonometric needlets. A FORTRAN implementation of this algorithm has been developed and extensively tested. The capabilities of the code are demonstrated using as examples the disturbing potential T, height anomaly ζ , gravity anomaly Δ g , gravity disturbance δ g , north-south deflection of the vertical ξ , east-west deflection of the vertical η , and the second radial derivative T_{rr} of the disturbing potential. After a pre-computational step that takes between 1 and 2 h per quantity, the current version of the software is capable of computing on a standard PC each of these quantities in the range from the surface of the Earth up to 544 km above that surface at speeds between 20,000 and 40,000 point evaluations per second, depending on the gravimetric quantity being evaluated, while the relative error does not exceed 10^{-6} and the memory (RAM) use is 9.3 GB.

Space Launch System Liftoff and Transition Aerodynamic Characterization in the NASA Langley 14- by 22-Foot Subsonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Pinier, Jeremy T.; Erickson, Gary E.; Paulson, John W.; Tomek, William G.; Bennett, David W.; Blevins, John A.

2015-01-01

A 1.75% scale force and moment model of the Space Launch System was tested in the NASA Langley Research Center 14- by 22-Foot Subsonic Wind Tunnel to quantify the aerodynamic forces that will be experienced by the launch vehicle during its liftoff and transition to ascent flight. The test consisted of two parts: the first was dedicated to measuring forces and moments for the entire range of angles of attack (0deg to 90deg) and roll angles (0 deg. to 360 deg.). The second was designed to measure the aerodynamic effects of the liftoff tower on the launch vehicle for ground winds from all azimuthal directions (0 deg. to 360 deg.), and vehicle liftoff height ratios from 0 to 0.94. This wind tunnel model also included a set of 154 surface static pressure ports. Details on the experimental setup, and results from both parts of testing are presented, along with a description of how the wind tunnel data was analyzed and post-processed in order to develop an aerodynamic database. Finally, lessons learned from experiencing significant dynamics in the mid-range angles of attack due to steady asymmetric vortex shedding are presented.

On the dependence of the domain of values of functionals of hypersonic aerodynamics on controls

NASA Astrophysics Data System (ADS)

Bilchenko, Grigory; Bilchenko, Nataly

2018-05-01

The properties of mathematical model of control of heat and mass transfer in laminar boundary layer on permeable cylindrical and spherical surfaces of the hypersonic aircraft are considered. Dependences of hypersonic aerodynamics functionals (the total heat flow and the total Newton friction force) on controls (the blowing into boundary layer, the temperature factor, the magnetic field) are investigated. The domains of allowed values of functionals of hypersonic aerodynamics are obtained. The results of the computational experiments are presented: the dependences of total heat flow on controls; the dependences of total Newton friction force on controls; the mutual dependences of functionals (as the domains of allowed values "Heat and Friction"); the dependences of blowing system power on controls. The influences of magnetic field and dissociation on the domain of "Heat and Friction" allowed values are studied. It is proved that for any fixed constant value of magnetic field the blowing system power is a symmetric function of constant dimensionless controls (the blowing into boundary layer and the temperature factor). It is shown that the obtained domain of allowed values of functionals of hypersonic aerodynamics depending on permissible range of controls may be used in engineering.

Experimental measurement of the aerodynamic charateristics of two-dimensional airfoils for an unmanned aerial vehicle

NASA Astrophysics Data System (ADS)

Velazquez, Luis; Nožička, Jiří; Vavřín, Jan

2012-04-01

This paper is part of the development of an airfoil for an unmanned aerial vehicle (UAV) with internal propulsion system; the investigation involves the analysis of the aerodynamic performance for the gliding condition of two-dimensional airfoil models which have been tested. This development is based on the modification of a selected airfoil from the NACA four digits family. The modification of this base airfoil was made in order to create a blowing outlet with the shape of a step on the suction surface since the UAV will have an internal propulsion system. This analysis involved obtaining the lift, drag and pitching moment coefficients experimentally for the situation where there is not flow through the blowing outlet, called the no blowing condition by means of wind tunnel tests. The methodology to obtain the forces experimentally was through an aerodynamic wire balance. Obtained results were compared with numerical results by means of computational fluid dynamics (CFD) from references and found in very good agreement. Finally, a selection of the airfoil with the best aerodynamic performance is done and proposed for further analysis including the blowing condition.

Flexible Multiplexed Surface Temperature Sensor

NASA Technical Reports Server (NTRS)

Daryabeigi, Kamran; Dillon-Townes, L. A.; Johnson, Preston B.; Ash, Robert L.

1995-01-01

Unitary array of sensors measures temperatures at points distributed over designated area on surface. Useful in measuring surface temperatures of aerodynamic models and thermally controlled objects. Made of combination of integrated-circuit microchips and film circuitry. Temperature-sensing chips scanned at speeds approaching 10 kHz. Operating range minus 40 degrees C to 120 degrees C. Flexibility of array conforms to curved surfaces. Multiplexer eliminates numerous monitoring cables. Control of acquisition and recording of data effected by connecting array to microcomputers via suitable interface circuitry.

Analysis of the surface heat balance over the world ocean

NASA Technical Reports Server (NTRS)

Esbensen, S. K.

1981-01-01

It is possible to estimate long term monthly mean latent and sensible heat fluxes over the ocean to within or approximately 20% relative accuracy of the bulk aerodynamic formulas, by using observations of the monthly mean surface wind speed and the monthly mean sea air temperature and humidity differences. It is possible to make an estimate of the fluxes on a month to month basis from monthly averaged surface data.

Aerodynamic Simulation of Runback Ice Accretion

NASA Technical Reports Server (NTRS)

Broeren, Andy P.; Whalen, Edward A.; Busch, Greg T.; Bragg, Michael B.

2010-01-01

This report presents the results of recent investigations into the aerodynamics of simulated runback ice accretion on airfoils. Aerodynamic tests were performed on a full-scale model using a high-fidelity, ice-casting simulation at near-flight Reynolds (Re) number. The ice-casting simulation was attached to the leading edge of a 72-in. (1828.8-mm ) chord NACA 23012 airfoil model. Aerodynamic performance tests were conducted at the ONERA F1 pressurized wind tunnel over a Reynolds number range of 4.7?10(exp 6) to 16.0?10(exp 6) and a Mach (M) number ran ge of 0.10 to 0.28. For Re = 16.0?10(exp 6) and M = 0.20, the simulated runback ice accretion on the airfoil decreased the maximum lift coe fficient from 1.82 to 1.51 and decreased the stalling angle of attack from 18.1deg to 15.0deg. The pitching-moment slope was also increased and the drag coefficient was increased by more than a factor of two. In general, the performance effects were insensitive to Reynolds numb er and Mach number changes over the range tested. Follow-on, subscale aerodynamic tests were conducted on a quarter-scale NACA 23012 model (18-in. (457.2-mm) chord) at Re = 1.8?10(exp 6) and M = 0.18, using low-fidelity, geometrically scaled simulations of the full-scale castin g. It was found that simple, two-dimensional simulations of the upper- and lower-surface runback ridges provided the best representation of the full-scale, high Reynolds number iced-airfoil aerodynamics, whereas higher-fidelity simulations resulted in larger performance degrada tions. The experimental results were used to define a new subclassification of spanwise ridge ice that distinguishes between short and tall ridges. This subclassification is based upon the flow field and resulting aerodynamic characteristics, regardless of the physical size of the ridge and the ice-accretion mechanism.

Aerodynamic Coefficients from Aeroballistic Range Testing of Deployed- and Stowed-SIAD SFDT Models

NASA Technical Reports Server (NTRS)

Wilder, Michael C.; Brown, Jeffrey D.; Bogdanoff, David W.; Yates, Leslie A.; Dyakonov, Artem A.; Clark, Ian G.; Grinstead, Jay H.

2017-01-01

This report documents a ballistic-range test campaign conducted in 2012 in order to estimate the aerodynamic stability characteristics of two configurations of the Supersonic Flight Dynamics Test (SFDT) vehicle prior to its initial flight in 2014. The SFDT vehicle was a test bed for demonstrating several new aerodynamic decelerator technologies then being developed under the Low-Density Supersonic Decelerator (LDSD) Project. Of particular interest here is the Supersonic Inflatable Aerodynamic Decelerator (SIAD), an inflatable attached torus used to increase the drag surface area of an entry vehicle during the supersonic portion of the entry trajectory. Two model configurations were tested in the ballistic range: one representing the SFDT vehicle prior to deployment of the SIAD, and the other representing the nominal shape with the SIAD inflated. Both models were fabricated from solid metal, and therefore, the effects of the flexibility of the inflatable decelerator were not considered. The test conditions were chosen to match, as close as possible, the Mach number, Reynolds number, and motion dynamics expected for the SFDT vehicle in flight, both with the SIAD stowed and deployed. For SFDT models with the SIAD stowed, 12 shots were performed covering a Mach number range of 3.2 to 3.7. For models representing the deployed SIAD, 37 shots were performed over a Mach number range of 2.0 to 3.8. Pitch oscillation amplitudes covered a range from 0.7 to 20.6 degrees RMS. Portions of this report (data analysis approach, aerodynamic modeling, and resulting aerodynamic coefficients) were originally published as an internal LDSD Project report [1] in 2012. In addition, this report provides a description of the test design approach, the test facility, and experimental procedures. Estimated non-linear aerodynamic coefficients, including pitch damping, for both model configurations are reported, and the shot-by-shot trajectory measurements, plotted in comparison with calculated trajectories based on the derived non-linear aerodynamic coefficients, are provided as appendices. Since the completion of these tests, two full-scale SFDT flights have been successfully conducted: one in June 2014 [2, 3], and one in June 2015 [3].

Large eddy simulation for aerodynamics: status and perspectives.

PubMed

Sagaut, Pierre; Deck, Sébastien

2009-07-28

The present paper provides an up-to-date survey of the use of large eddy simulation (LES) and sequels for engineering applications related to aerodynamics. Most recent landmark achievements are presented. Two categories of problem may be distinguished whether the location of separation is triggered by the geometry or not. In the first case, LES can be considered as a mature technique and recent hybrid Reynolds-averaged Navier-Stokes (RANS)-LES methods do not allow for a significant increase in terms of geometrical complexity and/or Reynolds number with respect to classical LES. When attached boundary layers have a significant impact on the global flow dynamics, the use of hybrid RANS-LES remains the principal strategy to reduce computational cost compared to LES. Another striking observation is that the level of validation is most of the time restricted to time-averaged global quantities, a detailed analysis of the flow unsteadiness being missing. Therefore, a clear need for detailed validation in the near future is identified. To this end, new issues, such as uncertainty and error quantification and modelling, will be of major importance. First results dealing with uncertainty modelling in unsteady turbulent flow simulation are presented.

Aerodynamic shape optimization using preconditioned conjugate gradient methods

NASA Technical Reports Server (NTRS)

Burgreen, Greg W.; Baysal, Oktay

1993-01-01

In an effort to further improve upon the latest advancements made in aerodynamic shape optimization procedures, a systematic study is performed to examine several current solution methodologies as applied to various aspects of the optimization procedure. It is demonstrated that preconditioned conjugate gradient-like methodologies dramatically decrease the computational efforts required for such procedures. The design problem investigated is the shape optimization of the upper and lower surfaces of an initially symmetric (NACA-012) airfoil in inviscid transonic flow and at zero degree angle-of-attack. The complete surface shape is represented using a Bezier-Bernstein polynomial. The present optimization method then automatically obtains supercritical airfoil shapes over a variety of freestream Mach numbers. Furthermore, the best optimization strategy examined resulted in a factor of 8 decrease in computational time as well as a factor of 4 decrease in memory over the most efficient strategies in current use.

An experimental study of the aerodynamics of a NACA 0012 airfoil with a simulated glaze ice accretion

NASA Technical Reports Server (NTRS)

Bragg, M. B.

1986-01-01

An experimental study was conducted in the Ohio State University subsonic wind tunnel to measure the detailed aerodynamic characteristics of an airfoil with a simulated glaze ice accretion. A NACA 0012 model with interchangeable leading edges and pressure taps every one percent chord was used. Surface pressure and wake data were taken on the airfoil clean, with forced transition and with a simulated glaze ice shape. Lift and drag penalties due to the ice shape were found and the surface pressure clearly showed that large separation bubbles were present. Both total pressure and split-film probes were used to measure velocity profiles, both for the clean model and for the model with a simulated ice accretion. A large region of flow separation was seen in the velocity profiles and was correlated to the pressure measurements. Clean airfoil data were found to compare well to existing airfoil analysis methods.

Aerodynamic characteristics of a Sparrow 3 missile model in the flow field of a generalized parent body at Mach 2.86

NASA Technical Reports Server (NTRS)

Stallings, R. L., Jr.

1984-01-01

Longitudinal aerodynamic characteristics of a Sparrow 3 wing control missile model were measured through a range of separation distances relative to a flat plate surface that represented the parent-body configuration. Measurements were obtained with and without two dimensional circular arc protuberances attached to the flat plate surface. The tests were conducted at a Mach number of 2.86 and a Reynolds number per meter of 6.56 million. The behavior of these longitudinal characteristics with varying separation distance in the flow field created by the flat plate and protuberance was generally as would be expected on the basis of flow field boundaries determined from the second order approximation of Friedrich. In general, varying roll angle from 0 deg to 45 deg caused no significant effect on the store separation characteristics.

The design, analysis, and testing of a low-budget wind-tunnel flutter model with active aerodynamic controls

NASA Technical Reports Server (NTRS)

Bolding, R. M.; Stearman, R. O.

1976-01-01

A low budget flutter model incorporating active aerodynamic controls for flutter suppression studies was designed as both an educational and research tool to study the interfering lifting surface flutter phenomenon in the form of a swept wing-tail configuration. A flutter suppression mechanism was demonstrated on a simple semirigid three-degree-of-freedom flutter model of this configuration employing an active stabilator control, and was then verified analytically using a doublet lattice lifting surface code and the model's measured mass, mode shapes, and frequencies in a flutter analysis. Preliminary studies were significantly encouraging to extend the analysis to the larger degree of freedom AFFDL wing-tail flutter model where additional analytical flutter suppression studies indicated significant gains in flutter margins could be achieved. The analytical and experimental design of a flutter suppression system for the AFFDL model is presented along with the results of a preliminary passive flutter test.

The Pressure Distribution over the Wings and Tail Surfaces of a PW-9 Pursuit Airplane in Flight

NASA Technical Reports Server (NTRS)

Rhode, Richard

1931-01-01

This report presents the results of an investigation to determine (1) the magnitude and distribution of aerodynamic loads over the wings and tail surfaces of a pursuit-type airplane in the maneuvers likely to impose critical loads on the various subassemblies of the airplane structure. (2) To study the phenomenon of center of pressure movement and normal force coefficient variation in accelerated flight, and (3) to measure the normal accelerations at the center of gravity, wing-tip, and tail, in order to determine the nature of the inertia forces acting simultaneously with the critical aerodynamic loads. The results obtained throw light on a number of important questions involving structural design. Some of the more interesting results are discussed in some detail, but in general the report is for the purpose of making this collection of airplane-load data obtained in flight available to those interested in airplane structures.

Unsteady aerodynamics of an oscillating cascade in a compressible flow field

NASA Technical Reports Server (NTRS)

Buffum, Daniel H.; Boldman, Donald R.; Fleeter, Sanford

1987-01-01

Fundamental experiments were performed in the NASA Lewis Transonic Oscillating Cascade Facility to investigate and quantify the unsteady aerodynamics of a cascade of biconvex airfoils executing torsion-mode oscillations at realistic reduced frequencies. Flush-mounted, high-response miniature pressure transducers were used to measure the unsteady airfoil surface pressures. The pressures were measured for three interblade phase angles at two inlet Mach numbers, 0.65 and 0.80, and two incidence angles, 0 and 7 deg. The time-variant pressures were analyzed by means of discrete Fourier transform techniques, and these unique data were then compared with predictions from a linearized unsteady cascade model. The experimental results indicate that the interblade phase angle had a major effect on the chordwise distributions of the airfoil surface unsteady pressure, and that reduced frequency, incidence angle, and Mach number had a somewhat less significant effect.

Non-monotonous dependence of the ocean surface drag coefficient on the hurricane wind speed due to the fragmentation of the ocean-atmosphere interface

NASA Astrophysics Data System (ADS)

Troitskaya, Yu. I.; Ermakova, O. S.; Kandaurov, A. A.; Kozlov, D. S.; Sergeev, D. A.; Zilitinkevich, S. S.

2017-11-01

Influence of the spray generation due to the fragmentation of the "bag-breakup" type on momentum exchange in the atmospheric boundary layer above the sea surface at hurricane winds was investigated on the basis of the analysis of the results of laboratory experiments. It was shown that aerodynamic drag is determined by the contribution of three factors: first, the drag of the "bag-breakup" canopies as obstacles; second, acceleration of the spray formed during fragmentation by the air flow; and the third factor is related to the stratification of the near-water atmospheric layer due to the presence of levitated water droplets. Combination of all three factors leads to a non-monotonous dependence of the aerodynamic drag coefficient on wind speed, which confirms the results of the field and laboratory measurements.

Aerodynamic configuration design using response surface methodology analysis

NASA Technical Reports Server (NTRS)

Engelund, Walter C.; Stanley, Douglas O.; Lepsch, Roger A.; Mcmillin, Mark M.; Unal, Resit

1993-01-01

An investigation has been conducted to determine a set of optimal design parameters for a single-stage-to-orbit reentry vehicle. Several configuration geometry parameters which had a large impact on the entry vehicle flying characteristics were selected as design variables: the fuselage fineness ratio, the nose to body length ratio, the nose camber value, the wing planform area scale factor, and the wing location. The optimal geometry parameter values were chosen using a response surface methodology (RSM) technique which allowed for a minimum dry weight configuration design that met a set of aerodynamic performance constraints on the landing speed, and on the subsonic, supersonic, and hypersonic trim and stability levels. The RSM technique utilized, specifically the central composite design method, is presented, along with the general vehicle conceptual design process. Results are presented for an optimized configuration along with several design trade cases.

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.

Enhancement of aerodynamic performance of a heaving airfoil using synthetic-jet based active flow control.

PubMed

Wang, Chenglei; Tang, Hui

2018-05-25

In this study, we explore the use of synthetic jet (SJ) in manipulating the vortices around a rigid heaving airfoil, so as to enhance its aerodynamic performance. The airfoil heaves at two fixed pitching angles, with the Strouhal number, reduced frequency and Reynolds number chosen as St  =  0.3, k  =  0.25 and Re  =  100, respectively, all falling in the ranges for natural flyers. As such, the vortex force plays a dominant role in determining the airfoil's aerodynamic performance. A pair of in-phase SJs is implemented on the airfoil's upper and lower surfaces, operating with the same strength but in opposite directions. Such a fluid-structure interaction problem is numerically solved using a lattice Boltzmann method based numerical framework. It is found that, as the airfoil heaves with zero pitching angle, its lift and drag can be improved concurrently when the SJ phase angle [Formula: see text] relative to the heave motion varies between [Formula: see text] and [Formula: see text]. But this concurrent improvement does not occur as the airfoil heaves with [Formula: see text] pitching angle. Detailed inspection of the vortex evolution and fluid stress over the airfoil surface reveals that, if at good timing, the suction and blowing strokes of the SJ pair can effectively delay or promote the shedding of leading edge vortices, and mitigate or even eliminate the generation of trailing edge vortices, so as to enhance the airfoil's aerodynamic performance. Based on these understandings, an intermittent operation of the SJ pair is then proposed to realize concurrent lift and drag improvement for the heaving airfoil with [Formula: see text] pitching angle.

Gust prediction via artificial hair sensor array and neural network

NASA Astrophysics Data System (ADS)

Pankonien, Alexander M.; Thapa Magar, Kaman S.; Beblo, Richard V.; Reich, Gregory W.

2017-04-01

Gust Load Alleviation (GLA) is an important aspect of flight dynamics and control that reduces structural loadings and enhances ride quality. In conventional GLA systems, the structural response to aerodynamic excitation informs the control scheme. A phase lag, imposed by inertia, between the excitation and the measurement inherently limits the effectiveness of these systems. Hence, direct measurement of the aerodynamic loading can eliminate this lag, providing valuable information for effective GLA system design. Distributed arrays of Artificial Hair Sensors (AHS) are ideal for surface flow measurements that can be used to predict other necessary parameters such as aerodynamic forces, moments, and turbulence. In previous work, the spatially distributed surface flow velocities obtained from an array of artificial hair sensors using a Single-State (or feedforward) Neural Network were found to be effective in estimating the steady aerodynamic parameters such as air speed, angle of attack, lift and moment coefficient. This paper extends the investigation of the same configuration to unsteady force and moment estimation, which is important for active GLA control design. Implementing a Recurrent Neural Network that includes previous-timestep sensor information, the hair sensor array is shown to be capable of capturing gust disturbances with a wide range of periods, reducing predictive error in lift and moment by 68% and 52% respectively. The L2 norms of the first layer of the weight matrices were compared showing a 23% emphasis on prior versus current information. The Recurrent architecture also improves robustness, exhibiting only a 30% increase in predictive error when undertrained as compared to a 170% increase by the Single-State NN. This diverse, localized information can thus be directly implemented into a control scheme that alleviates the gusts without waiting for a structural response or requiring user-intensive sensor calibration.

Simulation-Based Analysis of Reentry Dynamics for the Sharp Atmospheric Entry Vehicle

NASA Technical Reports Server (NTRS)

Tillier, Clemens Emmanuel

1998-01-01

This thesis describes the analysis of the reentry dynamics of a high-performance lifting atmospheric entry vehicle through numerical simulation tools. The vehicle, named SHARP, is currently being developed by the Thermal Protection Materials and Systems branch of NASA Ames Research Center, Moffett Field, California. The goal of this project is to provide insight into trajectory tradeoffs and vehicle dynamics using simulation tools that are powerful, flexible, user-friendly and inexpensive. Implemented Using MATLAB and SIMULINK, these tools are developed with an eye towards further use in the conceptual design of the SHARP vehicle's trajectory and flight control systems. A trajectory simulator is used to quantify the entry capabilities of the vehicle subject to various operational constraints. Using an aerodynamic database computed by NASA and a model of the earth, the simulator generates the vehicle trajectory in three-dimensional space based on aerodynamic angle inputs. Requirements for entry along the SHARP aerothermal performance constraint are evaluated for different control strategies. Effect of vehicle mass on entry parameters is investigated, and the cross range capability of the vehicle is evaluated. Trajectory results are presented and interpreted. A six degree of freedom simulator builds on the trajectory simulator and provides attitude simulation for future entry controls development. A Newtonian aerodynamic model including control surfaces and a mass model are developed. A visualization tool for interpreting simulation results is described. Control surfaces are roughly sized. A simple controller is developed to fly the vehicle along its aerothermal performance constraint using aerodynamic flaps for control. This end-to-end demonstration proves the suitability of the 6-DOF simulator for future flight control system development. Finally, issues surrounding real-time simulation with hardware in the loop are discussed.

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.).

The effect of tip shields on a horizontal tail surface

NASA Technical Reports Server (NTRS)

Dronin, Paul V; Ramsden, Earl I; Higgins, George J

1928-01-01

A series of experiments made in the wind tunnel of the Daniel Guggenheim School of Aeronautics, New York University, on the effect of tip shields on a horizontal tail surface are described and discussed. It was found that some aerodynamic gain can be obtained by the use of tip shields though it is considered doubtful whether their use would be practical.