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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 217 records · Page 12
All sourcesNASA NTRSDOE OSTI
NASA NTRS2005

Three-Dimensional Aeroelastic and Aerothermoelastic Behavior in Hypersonic Flow

The aeroelastic and aerothermoelastic behavior of three-dimensional configurations in hypersonic flow regime are studied. The aeroelastic behavior of a low aspect ratio wing, representative of a fin or control surface on a generic hypersonic vehicle, is examined using third order piston theory, Euler and Navier-Stokes aerodynamics. The sensitivity of the aeroelastic behavior generated using Euler and Navier-Stokes aerodynamics to parameters governing temporal accuracy is also examined. Also, a refined aerothermoelastic model, which incorporates the heat transfer between the fluid and structure using CFD generated aerodynamic heating, is used to examine the aerothermoelastic behavior of the low aspect ratio wing in the hypersonic regime. Finally, the hypersonic aeroelastic behavior of a generic hypersonic vehicle with a lifting-body type fuselage and canted fins is studied using piston theory and Euler aerodynamics for the range of 2.5 less than or equal to M less than or equal to 28, at altitudes ranging from 10,000 feet to 80,000 feet. This analysis includes a study on optimal mesh selection for use with Euler aerodynamics. In addition to the aeroelastic and aerothermoelastic results presented, three time domain flutter identification techniques are compared, namely the moving block approach, the least squares curve fitting method, and a system identification technique using an Auto-Regressive model of the aeroelastic system. In general, the three methods agree well. The system identification technique, however, provided quick damping and frequency estimations with minimal response record length, and therefore o ers significant reductions in computational cost. In the present case, the computational cost was reduced by 75%. The aeroelastic and aerothermoelastic results presented illustrate the applicability of the CFL3D code for the hypersonic flight regime.

McNamara, Jack J.↗
NASA NTRS2009

Measurement and Characterization of Space Shuttle Solid Rocket Motor Plume Acoustics

Lift-off acoustic environments generated by the future Ares I launch vehicle are assessed by the NASA Marshall Space Flight Center (MSFC) acoustics team using several prediction tools. This acoustic environment is directly caused by the Ares I First Stage booster, powered by the five-segment Reusable Solid Rocket Motor (RSRMV). The RSRMV is a larger-thrust derivative design from the currently used Space Shuttle solid rocket motor, the Reusable Solid Rocket Motor (RSRM). Lift-off acoustics is an integral part of the composite launch vibration environment affecting the Ares launch vehicle and must be assessed to help generate hardware qualification levels and ensure structural integrity of the vehicle during launch and lift-off. Available prediction tools that use free field noise source spectrums as a starting point for generation of lift-off acoustic environments are described in the monograph NASA SP-8072: "Acoustic Loads Generated by the Propulsion System." This monograph uses a reference database for free field noise source spectrums which consist of subscale rocket motor firings, oriented in horizontal static configurations. The phrase "subscale" is appropriate, since the thrust levels of rockets in the reference database are orders of magnitude lower than the current design thrust for the Ares launch family. Thus, extrapolation is needed to extend the various reference curves to match Ares-scale acoustic levels. This extrapolation process yields a subsequent amount of uncertainty added upon the acoustic environment predictions. As the Ares launch vehicle design schedule progresses, it is important to take every opportunity to lower prediction uncertainty and subsequently increase prediction accuracy. Never before in NASA s history has plume acoustics been measured for large scale solid rocket motors. Approximately twice a year, the RSRM prime vendor, ATK Launch Systems, static fires an assembled RSRM motor in a horizontal configuration at their test facility in Utah. The remaining RSRM static firings will take place on elevated terrain, with the nozzle exit plume being mostly undeflected and the landscape allowing placement of microphones within direct line of sight to the exhaust plume. These measurements will help assess the current extrapolation process by direct comparison between subscale and full scale solid rocket motor data.

Kenny, Jeremy↗
NASA NTRSDate not supplied

Training the Powered-Lift Evaluation Pilot

This poster describes a project to prepare pilots for a study assessing novel aircraft automation concepts for electric Vertical Takeoff and Landing (eVTOL) aircraft using NASA’s Vertical Motion Simulator (VMS). By exploring the operational and learning challenges related to transitioning between forward flight and vertical landing, we seek to establish baselines of pilot workload and aircraft handling qualities across varying atmospheric conditions and automation states. The simulated eVTOL design differentiates flight control allocations as a function of airspeed across four speed ranges as the vehicle transitions between fully thrust-borne lift and wing-borne lift. As speed increases, side stick controls command: translational ground speeds, vertical and lateral acceleration, vertical rate, vertical flight path angle, and bank angle. This novel approach to flight control allocation creates a significant learning challenge for pilots. Since initial eVTOL aircraft may have limitations on hover capabilities, automation and flight guidance cues also vary with airspeed to provide efficient landing profiles while still providing cues suitable for cruise flight. The NASA team prepared the study pilots to follow these flight guidance cues along curved Required Navigation Performance (RNP) approaches and along 6o and 12o glide paths to energy-efficient assistive-hover landing and goarounds. The pre-VMS preparation sought to prepare pilots from diverse levels of experience and background. To do this, NASA researchers designed and developed a fixed-based, large field-ofview simulator with terrain, structures, and air traffic. With one day of combined classroom learning and skill development in the fixedbase simulator, pilots were largely able to fly the simulated eVTOL in the VMS with sufficient mastery to provide handling quality assessments using the Cooper-Harper Handling Qualities Rating and workload assessments through the Bedford Workload Scale.

AAM↗
NASA NTRS1985

Harrier III-AV8B with a modern engine

The paper examines the application of a 26, 156-lb thrust engine (which is 30 percent more thrust than the short-lift dry rating on the F402-RR-406 Pegasus engine) to an AV-8B aircraft. This modern engine has 30 percent more thrust than the F402-RR-406 Pegasus engine with no increase in engine weight; the weight is assumed to be constant for both engines. The paper focuses on the benefit to the AV-8B in increasing short takeoff gross weight and vertical landing weight. A second comparison is made between the bigger engine with the existing thrust-versus-SFC curve, and a 10 percent lower SFC curve. The notion of a modern engine with 10 percent lower SFC and 30 percent more thrust for the same size and weight is presented as a hypothesis, and is not supported by discussing engine technology. There are no engine data or technical discussion as to how the modern engine is constructed or which manufacturer could build it. The performance estimates are generated by ACSYNT, a NASA developed computer code that is available to U.S. industry and has been reported on in a number of other publications.

Wilson, S. B., III↗
NASA NTRS1928

Wind Tunnel Tests on Autorotation and the "Flat Spin."

This report deals with the autorotational characteristics of certain differing wing systems as determined from wind tunnel tests made at the Langley Memorial Aeronautical Laboratory. The investigation was confined to autorotation about a fixed axis in the plane of symmetry and parallel to the wind direction. Analysis of the tests leads to the following conclusions: autorotation below 30 degree angle of attack is governed chiefly by wing profile, and above that angle by wing arrangement. The strip method of autorotation analysis gives uncertain results between maximum C subscript L and 35 degrees. The polar curve of a wing system, and to a lower degree of accuracy the polar of a complete airplane model are sufficient for direct determination of the limits of rotary instability, subject to strip method limitations. The results of the investigation indicate that in free flight a monoplane is incapable of flat spinning, whereas an unstaggered biplane has inherent flat-spinning tendencies. The difficulty of maintaining equilibrium in stalled flight is due primarily to rotary instability, a rapid change from stability to instability occurring as the angle of maximum lift is exceeded. (author)

Knight, Montgomery↗
NASA NTRS1931

An extended theory of thin airfoils and its application to the biplane problem

The report presents a new treatment, due essentially to von Karman, of the problem of the thin airfoil. The standard formulae for the angle of zero lift and zero moment are first developed and the analysis is then extended to give the effect of disturbing or interference velocities, corresponding to an arbitrary potential flow, which are superimposed on a normal rectilinear flow over the airfoil. An approximate method is presented for obtaining the velocities induced by a 2-dimensional airfoil at a point some distance away. In certain cases this method has considerable advantage over the simple "lifting line" procedure usually adopted. The interference effects for a 2-dimensional biplane are considered in the light of the previous analysis. The results of the earlier sections are then applied to the general problem of the interference effects for a 3-dimensional biplane, and formulae and charts are given which permit the characteristics of the individual wings of an arbitrary biplane without sweepback or dihedral to be calculated. In the final section the conclusions drawn from the application of the theory to a considerable number of special cases are discussed, and curves are given illustrating certain of these conclusions and serving as examples to indicate the nature of the agreement between the theory and experiment.

Millikan, Clark B↗
NASA NTRS1977

Aileron and Krueger nose flap effectiveness measured on an oblique wing

The effects that deflection of conventional ailerons have on the lateral/directional characteristics of an oblique wing mounted on top of a body of revolution were measured. The wing had an aspect ratio of 6 (based on the unswept wing span) and was tested at various sweep angles ranging from 0 deg to 60 deg. Krueger nose flaps mounted on drooped-nose flaps were investigated on the downstream wing panel (for sweep angles of 45 deg and 60 deg) in an attempt to create more symmetrical spanwise wing stalling at high lift coefficients. The tests were conducted over a Mach number range of 0.25 to 1.4 and an angle-of-attack range from -3 deg to 28 deg. With the oblique wing in the swept position, considerable difference in the aileron effectiveness was measured for positive or negative differential aileron deflections. With the wing swept 45 deg at a Mach number of 0.8 and above, an extension of the span of the Krueger nose flaps on the downstream wing panel did not improve the linearity of the pitching-, rolling-, and yawing-moment curves.

Hopkins, E. J.↗
NASA NTRS1972

Rene 95 brazed joint metallurgical program

This metallurgical program was specifically conducted for the establishment of material properties required for the design of the LF460 fan. The LF460 lift fan is an advanced 18:1 high thrust to weight single stage design. It has a turbine attached to the outer flowpath of the fan blade tip which minimizes the axial depth of the fan. Advanced lightweight attachment designs are employed in this concept to achieve minimum mass polar moments of inertia which are required for good aircraft flight response control. The design features which are unique to this advanced LF460 lift fan are the 0.010 inch thin Udimet 700 alloy integral tip turbine design, minimum weight braze attachment of the turbine to the fan blade, and the high strength and elevated temperature capability of the Rene'95 alloy for the fan blade. The data presented in this report show that the LF460 fan rotor design is feasible and that the design stresses and margins of safety were more than adequate. Prior to any production application, however, additional stress rupture/shear lap joints should be run in order to establish a firm 1200 F stress rupture curve for the CM50 braze metal.

Gay, C.↗
NASA NTRS1959

Effect of Convex Longitudinal Curvature on the Planing Characteristics of a Surface Without Dead Rise

A hydrodynamic investigation was made in Langley tank no. 1 of a planing surface which was curved longitudinally in the shape of a circular arc with the center of curvature above the model and had a beam of inches and a radius of curvature of 20 beams. The planing surface had length-beam ratio of 9 and an angle of dead rise of 0 deg. Wetted length, resistance, and trimming moment were determined for values of load coefficient C(sub Delta) from -4.2 to 63.9 and values of speed coefficient C(sub V) from 6 to 25. The effects of convexity were to increase the wetted length-beam ratio (for a given lift), to decrease the lift-drag ratio, to move the center of pressure forward, and ta increase the trim for maximum lift-drag ratio as compared with values for a flat surface. The effects were greatest at low trims and large drafts. The maximum negative lift coefficient C(sub L,b) obtainable with a ratio of the radius of curvature to the beam of 20 was -0.02. The effects of camber were greater in magnitude for convexity than for the same amount of concavity.

Mottard, Elmo J.↗
NASA NTRS1973

Evaluation of V/STOL research aircraft design

The evaluation and evolution of direct jet lift V/STOL transport aircraft designs are discussed. The V/STOL transport design selected as an example is a lift-fan design that was evaluated as a candidate configuration for a possible future V/STOL research transport. The paper includes discussion of potential advanced V/STOL landing approach profiles as key design requirements for V/STOL aircraft, description and experimental results of an integrated propulsion/control system designed to achieve desired advanced V/STOL near-terminal operating capabilities, and results from evaluating V/STOL designs on piloted moving-base simulators. This paper discusses use of the piloted moving-base simulator as a design tool for evolving satisfactory V/STOL stabilization and propulsion/control systems. Included are problems and solutions identified during simulation of simultaneous decelerating/descent steep curved landing approaches under instrument flight conditions. Simulation results are also compared to flight results obtained with the DO-31 V/STOL research transport.

Deckert, W. H.↗
NASA NTRS1984

Interactive aircraft flight control and aeroelastic stabilization

An analytical model has been used to study the problem of integrated control system/structural dynamic design. This model utilizes the traditional typical section lifting surface with a control surface attached in 2-D unsteady incompressible flow. The studies have as their purpose the examination of the use of the location of the elastic axis of the airfoil as a design variable in the search for an actively controlled configuration that has a specified flutter margin. Items of interest include the determination of flutter speed and divergence speed as functions of a nondimensional parameter, a sub c, that measures the location (in semi-chords) of the elastic axis with respect to the airfoil midchord. All other parameters are fixed. The behavior of velocity root locus curves with changes in a sub c is illustrated. Also shown is an example of the use of sensitivity derivatives to reposition poles of the open loop system. The behavior of the actively controlled or closed loop system is also discussed. A set of examples are presented to indicate how the shear center parameter, a sub c, affects the design of the controlled system.

Weisshaar, T. A.↗
NASA NTRS1986

Exact and approximate gas dynamics using the tangent gas

For the determination of aerodynamic characteristics such as lift, drag, and moment coefficients, it is crucial to compute the properties of steady flow past an airfoil. This investigation provides a set of flow dependent grid systems and initial flowfield guesses which substantially improve convergence rates when applied to the Euler equations for flows past an airfoil. The basic equations are examined, taking into account nonlinear equations which are difficult to solve. A good approximation to these equations under certain conditions can be obtained by introducing the so-called 'tangent gas approximation' considered by Woods (1961), in which the isentropic relation between rho and p is replaced by a tangent to a curve. Attention is given to the solution procedure, the analysis (direct) problem, and a comparison of the tangent gas solution with the converged Euler solution.

Daripa, P. K.↗
NASA NTRS2007

Drag Prediction for the DLR-F6 Wing/Body and DPW Wing using CFL3D and OVERFLOW Overset Mesh

A series of overset grids was generated in response to the 3rd AIAA CFD Drag Prediction Workshop (DPW-III) which preceded the 25th Applied Aerodynamics Conference in June 2006. DPW-III focused on accurate drag prediction for wing/body and wing-alone configurations. The grid series built for each configuration consists of a coarse, medium, fine, and extra-fine mesh. The medium mesh is first constructed using the current state of best practices for overset grid generation. The medium mesh is then coarsened and enhanced by applying a factor of 1.5 to each (I,J,K) dimension. The resulting set of parametrically equivalent grids increase in size by a factor of roughly 3.5 from one level to the next denser level. CFD simulations were performed on the overset grids using two different RANS flow solvers: CFL3D and OVERFLOW. The results were post-processed using Richardson extrapolation to approximate grid converged values of lift, drag, pitching moment, and angle-of-attack at the design condition. This technique appears to work well if the solution does not contain large regions of separated flow (similar to that seen n the DLR-F6 results) and appropriate grid densities are selected. The extra-fine grid data helped to establish asymptotic grid convergence for both the OVERFLOW FX2B wing/body results and the OVERFLOW DPW-W1/W2 wing-alone results. More CFL3D data is needed to establish grid convergence trends. The medium grid was utilized beyond the grid convergence study by running each configuration at several angles-of-attack so drag polars and lift/pitching moment curves could be evaluated. The alpha sweep results are used to compare data across configurations as well as across flow solvers. With the exception of the wing/body drag polar, the two codes compare well qualitatively showing consistent incremental trends and similar wing pressure comparisons.

Sclanfani, Anthony J.↗
NASA NTRS1977

Theory of viscous transonic flow over airfoils at high Reynolds number

This paper considers viscous flows with unseparated turbulent boundary layers over two-dimensional airfoils at transonic speeds. Conventional theoretical methods are based on boundary layer formulations which do not account for the effect of the curved wake and static pressure variations across the boundary layer in the trailing edge region. In this investigation an extended viscous theory is developed that accounts for both effects. The theory is based on a rational analysis of the strong turbulent interaction at airfoil trailing edges. The method of matched asymptotic expansions is employed to develop formal series solutions of the full Reynolds equations in the limit of Reynolds numbers tending to infinity. Procedures are developed for combining the local trailing edge solution with numerical methods for solving the full potential flow and boundary layer equations. Theoretical results indicate that conventional boundary layer methods account for only about 50% of the viscous effect on lift, the remaining contribution arising from wake curvature and normal pressure gradient effects.

Melnik, R. E.↗
NASA NTRS2005

A Program to Improve the Triangulated Surface Mesh Quality Along Aircraft Component Intersections

A computer program has been developed for improving the quality of unstructured triangulated surface meshes in the vicinity of component intersections. The method relies solely on point removal and edge swapping for improving the triangulations. It can be applied to any lifting surface component such as a wing, canard or horizontal tail component intersected with a fuselage, or it can be applied to a pylon that is intersected with a wing, fuselage or nacelle. The lifting surfaces or pylon are assumed to be aligned in the axial direction with closed trailing edges. The method currently maintains salient edges only at leading and trailing edges of the wing or pylon component. This method should work well for any shape of fuselage that is free of salient edges at the intersection. The method has been successfully demonstrated on a total of 125 different test cases that include both blunt and sharp wing leading edges. The code is targeted for use in the automated environment of numerical optimization where geometric perturbations to individual components can be critical to the aerodynamic performance of a vehicle. Histograms of triangle aspect ratios are reported to assess the quality of the triangles attached to the intersection curves before and after application of the program. Large improvements to the quality of the triangulations were obtained for the 125 test cases; the quality was sufficient for use with an automated tetrahedral mesh generation program that is used as part of an aerodynamic shape optimization method.

Cliff, Susan E.↗
NASA NTRS1977

Fluctuating loads on the flap surfaces of an externally-blown-flap configuration

An experimental program to obtain the fluctuating loads on the surfaces of a triple-slotted externally-blown-flap powered-lift STOL configuration was conducted. A large model of a wing/flap system and a TF-34 medium bypass ratio engine was investigated. Measurements of the fluctuating pressure, static pressure, and surface temperature resulting from the jet impingement were obtained at several locations on the surfaces of the second and third flaps. Fluctuating pressure data include overall level, power-spectral density (PSD), cross-correlation coefficient, coherency, and phase angle of the cross power-spectral density. These data indicate that more than one mechanism contributes to the fluctuating pressure levels on the flaps. In the immediate area above the intersection of the engine centerline and the flap, low frequency pressures dominate the overall fluctuating pressure levels. In other areas, such as below this intersection and outboard on the flaps, the PSD curve reaches a peak value at a Strouhal number ranging from 0.22 to 0.45.

Schoenster, J. A.↗
NASA NTRS2004

Aeroshell Design Techniques for Aerocapture Entry Vehicles

A major goal of NASA s In-Space Propulsion Program is to shorten trip times for scientific planetary missions. To meet this challenge arrival speeds will increase, requiring significant braking for orbit insertion, and thus increased deceleration propellant mass that may exceed launch lift capabilities. A technology called aerocapture has been developed to expand the mission potential of exploratory probes destined for planets with suitable atmospheres. Aerocapture inserts a probe into planetary orbit via a single pass through the atmosphere using the probe s aeroshell drag to reduce velocity. The benefit of an aerocapture maneuver is a large reduction in propellant mass that may result in smaller, less costly missions and reduced mission cruise times. The methodology used to design rigid aerocapture aeroshells will be presented with an emphasis on a new systems tool under development. Current methods for fast, efficient evaluations of structural systems for exploratory vehicles to planets and moons within our solar system have been under development within NASA having limited success. Many systems tools that have been attempted applied structural mass estimation techniques based on historical data and curve fitting techniques that are difficult and cumbersome to apply to new vehicle concepts and missions. The resulting vehicle aeroshell mass may be incorrectly estimated or have high margins included to account for uncertainty. This new tool will reduce the guesswork previously found in conceptual aeroshell mass estimations.

Dyke, R. Eric↗
NASA NTRS1993

Dynamic stall of an oscillating wing. Part 1: Evaluation of turbulence models

Unsteady flowfields of a two-dimensional oscillating wing are calculated using an implicit, finite-difference, Navier-Stokes numerical scheme using five widely used turbulence models. The objective of this study is to identify an appropriate turbulence model for accurate simulation of three-dimensional dynamic stall. Three unsteady flow conditions corresponding to attached flow, light-stall, and deep-stall of an oscillating wing experiment were chosen as test cases for computations. Results of unsteady airload hysteresis curves, harmonics of unsteady pressures, and instantaneous flow pictures are presented. Comparison of unsteady airloads with experiment show that all models are deficient in some sense and not a single model predicts all airloads consistently and in agreement with experiment for all flow conditions. For the attached flow condition, the Renormalization Group Theory (RNG), the Johnoson-King (J-K), and the Spalart-Allmaras (S-A) models have better performance. The Baldwin-Lomax (B-L) and the Baldwin-Barth (B-B) models fair poorly. At the light-stall condition, the results for the RNG, the J-K, and S-A models are in agreement with experiment for the upstroke but they all over predict the separation shown by the experiment and therefore have bigger hysteresis loops than experimental results. The B-B model results are also in good agreement for upstroke but have poor lift hysteresis for downstroke. It has superior drag and pitching-moment predictions. For deep-stall conditions, the airloads for the RNG, the B -B, and the S-A models have fair agreement with experiment, but the B-B model performed better at the extreme deep-stall condition. Overall, the RNG model provides significant improvement over the B-L model in all flow regimes with no additional computational cost. The Baldwin-Barth model is the most expensive of the models considered here, costing about 2.5 times that of the Baldwin-Lomax model. Finally, a brief discussion of the effects of grid density, time-step size, and numerical dissipation on the unsteady solutions are also presented.

Srinivasan, G. R.↗