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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 1,783 records · Page 99

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NASA NTRS1953

Preliminary Results of Stability and Control Investigation of the Bell X-5 Research Airplane

During the acceptance tests of the Bell X-5 airplane, measurements of the static stability and control characteristics and horizontal-tail loads were obtained by the NACA High-Speed Flight Research Station. The results of the stability and control measurements are presented in this paper. A change in sweep angle between 20 deg and 59 deg had a minor effect on the longitudinal trim, with a maximum change of about 2.5 deg in elevator deflection being required at a Mach number near 0.85; however, sweeping the wings produced a total stick-force change of about 40 pounds. At low Mach numbers there was a rapid increase in stability at high normal-force coefficients for both 20 0 and 1100 sweepback, whereas a condition of neutral stability existed for 58 0 sweepback at high normal-force coefficients. At Mach numbers near 0.8 there was an instability at normal-force coefficients above 0.5 for all sweep angles tested. In the low normal-force-coefficient range a high degree of stability resulted in high stick forces which limited the maximum load factors attainable in the demonstration flights to values under 5g for all sweep angles at a Mach number near 0.8 and an altitude of 12,000 feet. The aileron effectiveness at 200 sweepback was found to be low over the Mach number range tested.

Finch, Thomas W↗

NASA NTRS1976

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

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.

Nissim, E.↗

NASA NTRS1989

Structurally-Tailorable, Nonlinear, Snap-Through Spring

Abrupt change in load/deflection response controllable and predictable. Structurally-tailorable, nonlinear, snap-through spring (STNSTS) exhibits controllable and predictable abrupt change in load/deflection response and based upon known phenomenon of snap-through structural response. Composed of pin-connected two-bar linkage which depicts combined tension/compression springs. As load applied to STNSTS, stiffness is function of internal spring and bending stiffness of pin-connected bars. As load increases, bars deform laterally until they collapse and snap through. Has application in passively-tailored rotor-blade flap, pitch, and lag response, to improve aerodynamic performance and stability characteristics of rotors; in aerodynamically- and aeroelastically-tailored wing spars and ribs, to produce tailored deformation state for improved effectiveness in maneuvering, aerodynamic performance, and stability characteristics; and in energy absorbers for automobile bumpers and aircraft land

Starnes, James H., Jr.↗

NASA NTRS2000

Aeroelastic Response of Swept Aircraft Wings in a Compressible Flow Field

The present study addresses the subcritical aeroelastic response of swept wings, in various flight speed regimes, to arbitrary time-dependent external excitations. The methodology based on the concept of indicial functions is carried out in time and frequency domains. As a result of this approach, the proper unsteady aerodynamic loads necessary to study the subcritical aeroelastic response of the open/closed loop aeroelastic systems, and of flutter instability, respectively are obtained. Validation of the aeroelastic model is provided, and applications to subcritical aeroelastic response to blast pressure signatures are illustrated. In this context, an original representation of the aeroelastic response in the phase-space is displayed, and pertinent conclusions on the implications of a number of selected parameters of the system are outlined.

Marzocca, Piergiovanni↗

NASA NTRSDate not supplied

Design of a Manufacturing Demonstration Unit for a Composite Lunar Landing Gear Strut

The goal of the current work was to develop an analytical framework for design of composite struts using various levels of model fidelity. Rapid optimization trade studies were performed using low fidelity two-dimensional (2D) axisymmetric models with smeared composite properties. The optimum 2D model was compared with three-dimensional (3D) models with higher levels of fidelity in material property representation. Good agreement was found with all models. The buckling performance of the highest fidelity 3D model was found to be satisfactory for the intended loading conditions. An additional goal of this work was to perform an initial assessment of using automated fiber placement (AFP) and other advanced manufacturing methods to explore their feasibility for the fabrication of struts for lunar landers, strut-braced wings, and other aerospace components. The designs produced during this study are intended to be used to develop manufacturing demonstration units (MDU) that can be fabricated at the Integrated Structural Assembly of Advanced Composites (ISAAC) facility at Langley Research Center (LaRC) and tested in lab facilities at LaRC.

carbon-epoxy↗

NASA NTRSDate not supplied

Design of a Manufacturing Demonstration Unit for a Composite Lunar Landing Gear Strut

carbon-epoxy↗

DOE OSTI2025

Vitrification Testing of HLW with High Phosphate

Projections of the number of high level waste (HLW) canisters to be produced in the Hanford Tank Waste Treatment and Immobilization Plant (WTP) (e.g., [1]) are based upon the inventory of the tank wastes, the anticipated performance of the sludge treatment processes, and current understanding of the capability of the borosilicate glass waste form. The WTP HLW melter design, unlike earlier Department of Energy (DOE) melter designs, incorporates a glass bubbler system. The bubblers create active glass pool mixing and thereby improve heat and mass transfer and glass melting rate. The WTP HLW melters each have a glass surface area of 3.75 m 2 and depth of ~1.1 m. The two melters in the HLW facility together are designed to produce up to 7.5 MT of glass per day at 100% availability. Further increases in HLW waste processing rates can potentially be achieved by optimization of the feed and glass formulations, increasing the melter operating temperature above 1150⁰C, and by increasing the waste loading in the glass product. Increasing the waste loading also has the added benefit of decreasing the number of canisters for storage.

12 MANAGEMENT OF RADIOACTIVE AND NON-RADIOACTIVE W↗

NASA NTRS2015

Optimization of an Aeroservoelastic Wing with Distributed Multiple Control Surfaces

This paper considers the aeroelastic optimization of a subsonic transport wingbox under a variety of static and dynamic aeroelastic constraints. Three types of design variables are utilized: structural variables (skin thickness, stiffener details), the quasi-steady deflection scheduling of a series of control surfaces distributed along the trailing edge for maneuver load alleviation and trim attainment, and the design details of an LQR controller, which commands oscillatory hinge moments into those same control surfaces. Optimization problems are solved where a closed loop flutter constraint is forced to satisfy the required flight margin, and mass reduction benefits are realized by relaxing the open loop flutter requirements.

Stanford, Bret K.↗

DOE OSTI2014

Dynamic load balance scheme for the DSMC algorithm

The direct simulation Monte Carlo (DSMC) algorithm, devised by Bird, has been used over a wide range of various rarified flow problems in the past 40 years. While the DSMC is suitable for the parallel implementation on powerful multi-processor architecture, it also introduces a large load imbalance across the processor array, even for small examples. The load imposed on a processor by a DSMC calculation is determined to a large extent by the total of simulator particles upon it. Since most flows are impulsively started with initial distribution of particles which is surely quite different from the steady state, the total of simulator particles will change dramatically. The load balance based upon an initial distribution of particles will break down as the steady state of flow is reached. The load imbalance and huge computational cost of DSMC has limited its application to rarefied or simple transitional flows. In this paper, by taking advantage of METIS, a software for partitioning unstructured graphs, and taking the total of simulator particles in each cell as a weight information, the repartitioning based upon the principle that each processor handles approximately the equal total of simulator particles has been achieved. The computation must pause several times to renew the total of simulator particles in each processor and repartition the whole domain again. Thus the load balance across the processors array holds in the duration of computation. The parallel efficiency can be improved effectively. The benchmark solution of a cylinder submerged in hypersonic flow has been simulated numerically. Besides, hypersonic flow past around a complex wing-body configuration has also been simulated. The results have displayed that, for both of cases, the computational time can be reduced by about 50%.

71 CLASSICAL AND QUANTUM MECHANICS, GENERAL PHYSIC↗

DOE OSTI2015

Experimental analysis of energy harvesting from self-induced flutter of a composite beam

Previous attempts to harvest energy from aeroelastic vibrations have been based on attaching a beam to a moving wing or structure. Here, we exploit self-excited oscillations of a fluttering composite beam to harvest energy using piezoelectric transduction. Details of the beam properties and experimental setup are presented. The effects of preset angle of attack, wind speed, and load resistance on the levels of harvested power are determined. The results point to a complex relation between the aerodynamic loading and its impact on the static deflection and amplitudes of the limit cycle oscillations on one hand and the load resistance and level of power harvested on the other hand.

71 CLASSICAL AND QUANTUM MECHANICS, GENERAL PHYSIC↗

NASA NTRS1975

Results of investigation IA110 on a 0.015-scale integrated configuration of the space shuttle vehicle in the arc 9 x 7 supersonic wind tunnel using models 67-TS and 49-0

An 0.015-scale space shuttle vehicle model was tested to investigate Orbiter wind bending, elevon panel loads, and elevon effectiveness. Mach numbers from 1.5 through 2.5 were investigated. Angles of attack and sideslip were varied from -8 degrees through +8 degrees. Post test analysis of raw wind tunnel data indicated a zero shift had occured in the wing bending and torsional gages during the test. The mechanism by which this shift occurred was not determined. Therefore, all the wind root bending and torsional moment data is suspect.

Chee, E.↗

NASA NTRS1975

Application of local linearization and the transonic equivalence rule to the flow about slender analytic bodies at Mach numbers near 1.0

The local linearization method for axisymmetric flow is combined with the transonic equivalence rule to calculate pressure distribution on slender bodies at free-stream Mach numbers from .8 to 1.2. This is an approximate solution to the transonic flow problem which yields results applicable during the preliminary design stages of a configuration development. The method can be used to determine the aerodynamic loads on parabolic arc bodies having either circular or elliptical cross sections. It is particularly useful in predicting pressure distributions and normal force distributions along the body at small angles of attack. The equations discussed may be extended to include wing-body combinations.

Tyson, R. W.↗

NASA NTRS1981

Supersonic cruise research aircraft structural studies: Methods and results

NASA Langley Research Center SCAR in-house structural studies are reviewed. In methods development, advances include a new system of integrated computer programs called ISSYS, progress in determining aerodynamic loads and aerodynamically induced structural loads (including those due to gusts), flutter optimization for composite and metal airframe configurations using refined and simplified mathematical models, and synthesis of active controls. Results given address several aspects of various SCR configurations. These results include flutter penalties on composite wing, flutter suppression using active controls, roll control effectiveness, wing tip ground clearance, tail size effect on flutter, engine weight and mass distribution influence on flutter, and strength and flutter optimization of new configurations. The ISSYS system of integrated programs performed well in all the applications illustrated by the results, the diversity of which attests to ISSYS' versatility.

Sobieszczanski-Sobieski, J.↗

NASA NTRS1991

Technology needs for high-speed rotorcraft

A study to determine the technology development required for high-speed rotorcraft development was conducted. The study begins with an initial assessment of six concepts capable of flight at, or greater than 450 knots with helicopter-like hover efficiency (disk loading less than 50 pfs). These concepts were sized and evaluated based on measures of effectiveness and operational considerations. Additionally, an initial assessment of the impact of technology advances on the vehicles attributes was made. From these initial concepts a tilt wing and rotor/wing concepts were selected for further evaluation. A more detailed examination of conversion and technology trade studies were conducted on these two vehicles, each sized for a different mission.

Rutherford, John↗

NASA NTRS1992

Phoenix: Preliminary design of a high speed civil transport

The goal of the Phoenix Design Project was to develop a second generation high speed civil transport (HSCT) that will meet the needs of the traveler and airline industry beginning in the 21st century. The primary emphasis of the HSCT is to take advantage of the growing needs of the Pacific Basin and the passengers who are involved in that growth. A passenger load of 150 persons, a mission range of 5150 nautical miles, and a cruise speed of Mach 2.5 constitutes the primary design points of this HSCT. The design concept is made possible with the use of a well designed double delta wing and four mixed flow engines. Passenger comfort, compatibility with existing airport infrastructure, and cost competitive with current subsonic aircraft make the Phoenix a viable aircraft for the future.

Aguilar, Joseph↗

NASA NTRS2001

A Closed-Loop Optimal Neural-Network Controller to Optimize Rotorcraft Aeromechanical Behaviour: Theory and Methodology - Volume 1

Given the predicted growth in air transportation, the potential exists for significant market niches for rotary wing subsonic vehicles. Technological advances which optimise rotorcraft aeromechanical behaviour can contribute significantly to both their commercial and military development, acceptance, and sales. Examples of the optimisation of rotorcraft aeromechanical behaviour which are of interest include the minimisation of vibration and/or loads. The reduction of rotorcraft vibration and loads is an important means to extend the useful life of the vehicle and to improve its ride quality. Although vibration reduction can be accomplished by using passive dampers and/or tuned masses, active closed-loop control has the potential to reduce vibration and loads throughout a.wider flight regime whilst requiring less additional weight to the aircraft man that obtained by using passive methads. It is ernphasised that the analysis described herein is applicable to all those rotorcraft aeromechanical behaviour optimisation problems for which the relationship between the harmonic control vector and the measurement vector can be adequately described by a neural-network model.

Leyland, Jane Anne↗

NASA NTRS1947

Spin and Recovery Characteristics of the Curtiss-Wright XP-87 Airplane

The spin and recovery characteristics of the Curtiss-Wright XP-87 airplane, as well as the spin-recovery parachute requirements, the control forces that would be encountered in the spin, and the best method for the crew to attempt an emergency escape, are presented in this report. The characteristics were estimated rather than determined by model tests because the XP-87 dimensional and mass characteristics were considered to be noncritical and because data were available from model tests of several similar airplanes. The study indicated that the recovery characteristics of the airplane will be satisfactory for all loadings if the controls are reversed fully and rapidly. The control forces, however, will probably be beyond the capabilities of the pilot unless some additional balance or a booster is used. A 6-foot tail parachute or a 3.5-foot wing-tip parachute with a drag coefficient of 0.7 will be a satisfactory, emergency spin-recovery device for spin demonstrations. If it is necessary for the crew to abandon the spinning airplane, they should leave from the outboard side of the cockpit.

Berman, Theodore↗

NASA NTRS2011

Reliability Based Design for a Raked Wing Tip of an Airframe

A reliability-based optimization methodology has been developed to design the raked wing tip of the Boeing 767-400 extended range airliner made of composite and metallic materials. Design is formulated for an accepted level of risk or reliability. The design variables, weight and the constraints became functions of reliability. Uncertainties in the load, strength and the material properties, as well as the design variables, were modeled as random parameters with specified distributions, like normal, Weibull or Gumbel functions. The objective function and constraint, or a failure mode, became derived functions of the risk-level. Solution to the problem produced the optimum design with weight, variables and constraints as a function of the risk-level. Optimum weight versus reliability traced out an inverted-S shaped graph. The center of the graph corresponded to a 50 percent probability of success, or one failure in two samples. Under some assumptions, this design would be quite close to the deterministic optimum solution. The weight increased when reliability exceeded 50 percent, and decreased when the reliability was compromised. A design could be selected depending on the level of risk acceptable to a situation. The optimization process achieved up to a 20-percent reduction in weight over traditional design.

Patnaik, Surya N.↗