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At least 325 records · Page 18
All sourcesNASA NTRSDOE OSTI
NASA NTRS1994

Sub-ply level scaling approach investigated for graphite-epoxy composite beam columns

Scale model graphite-epoxy composite specimens were fabricated using the 'sub-ply level' approach and tested as beam-columns under an eccentric axial load to determine the effect of specimen size on flexural response and failure. In the current research project, although the fiber diameters are not scaled, the thickness of the pre-preg material itself has been scaled by adjusting the number of fibers through the thickness of a single ply. Three different grades of graphite-epoxy composite material (AS4/3502) were obtained from Hercules, Inc., in which the number of fibers through the thickness of a single ply was reduced (Grade 190 with 12 to 16 fibers, Grade 95 with 6 to 8 fibers, and Grade 48 with 3 to 4 fibers). Thus, using the sub-ply level approach, a baseline eight ply quasi-isotropic laminate could be fabricated using either the Grade 48 or Grade 95 material and the corresponding full-scale laminate would be constructed from Grade 95 or standard Grade 190 material, respectively. Note that in the sub-ply level approach, the number of ply interfaces is constant for the baseline and full-scale laminates. This is not true for the ply level and sublaminate level scaled specimens. The three grades of graphite-epoxy composite material were used to fabricate scale model beam-column specimens with in-plane dimensions of 0.5*n x 5.75*n, where n=1,2,4 corresponsing to 1/4, 1/2, and full-scale factors. Angle ply, cross ply, and quasi-isotropic laminate stacking sequences were chosen for the investigation and the test matrices for each laminate type are given. Specimens in each laminate family with the same in-plane dimensions but different thicknesses were tested to isolate the influence of the thickness dimension on the flexural response and failure. Also, specific lay-ups were chosen with blocked plies and dispersed plies for each laminate type. Specimens were subjected to an eccentric axial load until failure. The load offset was introduced through a set of hinges which were attached to the platens of a standard load test machine. Three sets of geometrically scaled hinges were used to ensure that scaled loading conditions were applied. This loading condition was chosen because it promotes large flexural deformations and specimens fail at the center of the beam, away from the grip supports. Five channels of data including applied vertical load, end shortening displacement, strain from gages applied back-to-back at the midspan of the beam, and rotation of the hinge from a bubble inclinometer were recorded for each specimen. The beam-column test configuration was used previously to study size effects in ply level scaled composite specimens of the same material system, sizes, and stacking sequences. Thus, a direct comparison between the two scaling approaches is possible. Ply level scaled beam-columns with angle ply, cross ply, and quasi-isotropic lay-ups exhibited no size dependencies in the flexural response, but significant size effects in strength. The reduction in strength with increasing specimen size was not predicted successfully by analysis techniques. It is anticipated that results from this investigation will lead to a better understanding of the strength scale effect in composite structures.

Jackson, Karen E.↗
NASA NTRS1944

A Theoretical Investigation of Longitudinal Stability of Airplanes with Free Controls Including Effect of Friction in Control System

The relation between the elevator hinge moment parameters and the control forces for changes in forward speed and in maneuvers is shown for several values of static stability and elevator mass balance. The stability of the short period oscillations is shown as a series of boundaries giving the limits of the stable regions in terms of the elevator hinge moment parameters. The effects of static stability, elevator moment of inertia, elevator mass unbalance, and airplane density are also considered. Dynamic instability is likely to occur if there is mass unbalance of the elevator control system combined with a small restoring tendency (high aerodynamic balance). This instability can be prevented by a rearrangement of the unbalancing weights which, however, involves an increase of the amount of weight necessary. It can also be prevented by the addition of viscous friction to the elevator control system provided the airplane center of gravity is not behind a certain critical position. For high values of the density parameter, which correspond to high altitudes of flight, the addition of moderate amounts of viscous friction may be destabilizing even when the airplane is statically stable. In this case, increasing the viscous friction makes the oscillation stable again. The condition in which viscous friction causes dynamic instability of a statically stable airplane is limited to a definite range of hinge moment parameters. It is shown that, when viscous friction causes increasing oscillations, solid friction will produce steady oscillations having an amplitude proportional to the amount of friction.

Greenberg, Harry↗
NASA NTRS1959

Full-Scale Wind-Tunnel Investigation of a Jet Flap in Conjunction with a Plain Flap with Blowing Boundary-Layer Control on a 35 deg Sweptback-Wing Airplane

Previous investigations have shown that increased blowing at the hinge-line radius of a plain flap will give flap lift increases above that realized with boundary-layer control. Other experiments and theory have shown that blowing from a wing trailing edge, through the jet flap effect, produced lift increases. The present investigation was made to determine whether blowing simultaneously at the hinge-line radius and trailing edge would be more effective than blowing separately at either location. The tests were made at a Reynolds number of 4.5 x 10(exp 6) with a 35 deg sweptback-wing airplane. For this report, only the lift data are presented. Of the three flap blowing arrangements tested, blowing distributed between the trailing edge and the hinge-line radius of a plain flap was found to be superior to blowing at either location separately at the plain flap deflections of interest. Comparison of estimated and experimental jet flap effectiveness was fair.

Aoyagi, Kiyoshi↗
NASA NTRS1947

Tests of a Horizontal-Tail Model through the Transonic Speed Range by the NACA Wing-Flow Method

A 1/12-scale model of a horizontal tail of a fighter airplane was tested through the transonic speeds in the high-speed flow over an airplane wing, the surface of which served as a reflection plane for the model. Measurements of lift, elevator-hinge moment, angle of attack, and elevator angle were made in the Mach number range from 0.75 to 1.04 for elevator deflections ranging from 10 degrees to minus 10 degrees, and for angles of attack of minus 1.2 degrees, 0.4 degrees, and 3.4 degrees. The equipment used to measure the hinge moments of the model proved to be unsatisfactory, and for this reason the hinge-moment data are considered to be only qualitative.

Adams, Richard E.↗
NASA NTRS1982

Rigid collapsible dish structure

A collapsible dish structure composed of a plurality of rows of rigid radial petal assemblies concentric with the axis of the dish. The petal assemblies consist of a center petal and two side petals, the center petal hinged on an axis tangent to a circle concentric with the axis of the dish and the side petals hinged to the center petal at their mating edge. The center petal is foldable inwardly and the side petals rotate about their hinges such that the collapsed dish structure occupies a much smaller volume than the deployed dish. Means of controlling the shape of the dish to compensate for differential expansion of the deployed dish are also provided.

Palmer, William B.↗
NASA NTRS2014

Artificial Intelligence Based Control Power Optimization on Tailless Aircraft

Traditional methods of control allocation optimization have shown difficulties in exploiting the full potential of controlling large arrays of control devices on innovative air vehicles. Artificial neutral networks are inspired by biological nervous systems and neurocomputing has successfully been applied to a variety of complex optimization problems. This project investigates the potential of applying neurocomputing to the control allocation optimization problem of Hybrid Wing Body (HWB) aircraft concepts to minimize control power, hinge moments, and actuator forces, while keeping system weights within acceptable limits. The main objective of this project is to develop a proof-of-concept process suitable to demonstrate the potential of using neurocomputing for optimizing actuation power for aircraft featuring multiple independently actuated control surfaces. A Nastran aeroservoelastic finite element model is used to generate a learning database of hinge moment and actuation power characteristics for an array of flight conditions and control surface deflections. An artificial neural network incorporating a genetic algorithm then uses this training data to perform control allocation optimization for the investigated aircraft configuration. The phase I project showed that optimization results for the sum of required hinge moments are improved by more than 12% over the best Nastran solution by using the neural network optimization process.

Gern, Frank↗
NASA NTRS2016

System Assessment of a High Power 3-U CubeSat

The Advanced eLectrical Bus (ALBus) CubeSat project is a technology demonstration mission of a 3-UCubeSat with an advanced, digitally controlled electrical power system capability and novel use of Shape Memory Alloy (SMA) technology for reliable deployable solar array mechanisms. The objective of the project is to, through an on orbit demonstration, advance the state of power management and distribution (PMAD) capabilities to enable future missions requiring higher power, flexible and reliable power systems. The goals of the mission include demonstration of: 100 Watt distribution to a target electrical load, efficient battery charging in the orbital environment, flexible power system distribution interfaces, adaptation of power system control on orbit, and reliable deployment of solar arrays and antennas utilizing re-settable SMA mechanisms. The power distribution function of the ALBus PMAD system is unique in the total power to target load capability of 100 W, the flexibility to support centralized or point-to-load regulation and ability to respond to fast transient power requirements. Power will be distributed from batteries at 14.8 V, 6.5 A to provide 100 W of power directly to a load. The deployable solar arrays utilize NASA Glenn Research Center superelastic and activated Nitinol(Nickel-Titanium alloy) Shape Memory Alloy (SMA) technology for hinges and a retention and release mechanism. The deployable solar array hinge design features utilization of the SMA material properties for dual purpose. The hinge uses the shape memory properties of the SMA to provide the spring force to deploy the arrays. The electrical conductivity properties of the SMA also enables the design to provide clean conduits for power transfer from the deployable arrays to the power management system. This eliminates the need for electrical harnesses between the arrays and the PMAD system in the ALBus system design. The uniqueness of the SMA retention and release mechanism design is the ability to reset the mechanism, allowing functional tests of the mechanisms prior to flight with no degradation of performance. The project is currently in preparation at the NASA Glenn Research Center for a launch in late calendar year of 2017. The 100 Watt power distribution and dual purpose, re-settable SMA mechanisms introduced several system level challenges due to the physical constraints in volume, mass and surface area of 3-U CubeSats. Several trade studies and design cycles have been completed to develop a system which supports the project objectives. This paper is a report on the results of the system level trade studies and assessments. The results include assessment of options for thermal control of 100 Watts of power dissipation, data from system analyses and engineering development tests, limitations of the 3-U system and extensibility to larger scale CubeSat missions.

CubeSat↗
NASA NTRSDate not supplied

Advanced eLectrical Bus (ALBus) CubeSat: From Build to Flight

The Advanced eLectrical Bus (ALBus) CubeSat project is a technology demonstration mission of a 3-U CubeSat with an advanced, digitally controlled electrical power system capability and novel use of Shape Memory Alloy (SMA) technology for reliable deployable solar array mechanisms. The objective of the project was to advance the state of power management and distribution (PMAD) capabilities to enable future missions requiring flexible and reliable power systems with higher output power capabilities. The goals of the mission were to demonstrate 100W distribution to a target electrical load, respond to continuous and fast transient power requirements, and show reliable deployment of solar arrays and antennas utilizing re-settable SMA mechanisms. The power distribution function of the ALBus PMAD system is unique in the total power to target load capability. Power is distributed from batteries to provide 100W of power directly to a resistive load. The deployable solar arrays utilize NASA Glenn Research Center’s Nickel-Titanium-Palladium-Platinum high-temperature SMAs for the retention and release mechanism, and a super-elastic binary NiTi alloy for the hinge component. The deployable solar array hinge design features the utilization of SMA material properties for a dual purpose. First, the hinge uses the super-elastic SMA properties to provide the spring force to deploy the arrays. Second, the electrical conductivity properties of the SMA enables the design to provide clean conduits for power transfer from the deployable arrays to the power management system. This eliminates the need for electrical harnesses between the arrays and the PMAD system in the ALBus system design. The uniqueness of the SMA retention and release mechanism design is the ability to reset the mechanism and reuse the same hardware. This allow functional tests of the mechanisms before flight with no degradation of performance and flying the hardware that is tested. The project launched as part of the CubeSat Launch Initiative Educational Launch of Nanosatellites (ELaNa) XIX mission on Rocket Lab’s Electron in December 2018. This paper summarizes the final launched design of ALBus and lessons learned from the technology and overall project from development through flight.

Deboshri Sadhukhan↗
NASA NTRS1966

Calculated blade response at high tip-speed ratios

This paper presents the initial results of a digital computer study of helicopter rotor-blade-motion stability for a broad range of forward speeds encompassing proposed conventional and compound helicopter designs. The analysis treats the general case of nonlinear, coupled flapping and lagging motions of hinged rigid blades. The results of the study to date indicate that blade-motion stability boundaries cannot be specified for a given blade design in terms of a fixed value of rotor tip-speed ratio. The stability boundaries can shift significantly within the desired operational speed range, with the stability limits depending upon the rotor loading, the mag­nitude of the disturbance encountered by the blade, and the blade position at the instant the external disturbance is encountered. The sample methods suggested for improving blade-motion stability include increasing the effec­tive hinge spring restraint by incorporating pitch-flap coupling or by using hingeless or teetering rotor systems. Reducing rotor loading also has a beneficial effect on the blade-motion stability. The results indi­cate that methods used to deal with the blade-motion stability problem can be expected to diminish the overall vibration levels of the helicopter rotor system across the entire design speed range.

TIP SPEED↗
NASA NTRS1947

Investigation of a Spoiler-Type Lateral Control System on a Wing with Full-Span Flaps in the Langley 19-Foot Pressure Tunnel

Tests of a partial-span model of a large bomber-type air1ane were conducted to determine the. aerodynamic characteristics of the wing equipped with full-span flaps and a retractable spoiler end aileron lateral control system. The arrangement consisted of (1) a double slotted flap extending over aproximate1y 86 percent of the wing semispan, (2) a 20-percent constant-percentage-chord aileron extending from the outboard end of the flap to the wing tip, and (3) a retractable spoiler, located at the 65-percent wing-chord station and extending from approximately 63 percent of the wing semispan to the wing tip. In addition, tests were made of a wing vent (of 1 and 2 percent of the wing chord located directly behind the spoiler), perforations in the spoiler, a blot or cut-out along the lower edge of the spoiler and spoilers of various spans. With full-span flaps deflected and with the 2-percent vent open or closed the initial stalling of the wing occurred at the tips, but with the vents closed there probably would be no appreciable loss in lateral control until maximum lift was reached. The l-percent vent increased the rolling effectiveness of the spoiler at small spoi1er deflections, particularly at high angles of attack with flaps deflected. With flaps deflected the 2-percent vent caused a large reduction in both the wing lift and rolling effectiveness of the spoiler at large angles of attack. However, at small angle of attack the 2-percent vent increased the rolling effectiveness of the spoiler at small spoiler deflections. The simultaneous operation of the spoiler and vent (in contrast to a vent fixed in the wing) would result in a large increase in the effectiveness of the spoiler and would avoid any loss in wing lift as in a fixed vent arrangement. The tests of the spoiler modifications revealed that (1) the spoiler perforations reduced the rolling-moment and yawing-moment coefficients but caused the spoiler hinge-moment coefficients to become more positive; (2) the spoiler slot had no notable effect on the rolling-moment and yawing-moment characteristics but produced a positive increase in the spoiler hinge-moment coefficients at large spoiler deflections; (3) the effects produced by the individual modifications were additive when the various modifications were combined. In general, progressively decreasing the spoiler span by removing the segments from the inboard end of the spoiler caused a decrease in rolling effectiveness approximately proportional to the span of the segment.

WING FLAPS, SLOTTED & AILERONS↗
NASA NTRS1942

Application of Balancing Tabs to Ailerons

Analysis was made to determine characteristics required of a balancing-tab system for ailerons in order to reduce aileron stick forces to any desired magnitude. Series of calculations based on section data were made to determine balancing-tab systems of various chord tabs and ailerons that will give, for a particular airplane, zero rate of aileron hinge moment with aileron deflection and yet will produce same maximum rate of roll as a plain unbalanced 15-percent chord aileron of same span. Effects of rolling velocity and of forces in tab link on aileron hinge moments have been included.

TABS, AILERON↗
NASA NTRS1944

Effect of Changes in Aileron Rigging on the Stick Forces of a High-Speed Fighter Airplane

The effects of changes in aileron rigging between 2 deg up and 2 deg down on the stick forces were determined from wind-tunnel data for a finite-span wing model. These effects were investigated for ailerons deflecting equally in both directions and linearly with stick deflection. Data were analyzed for a Frise, a sealed internally balanced, and a beveled-trailing-edge aileron. The results of the analysis showed that only ailerons having linear hinge-moment characteristics are unaffected by changes in rigging and indicated that ailerons having decidedly nonlinear hinge-moment-coefficient curves, particularly for deflections near 0 deg, are very sensitive to changes in rigging.

AILERONS, BEVELED↗
NASA NTRS2006

Verification of Numerical Solutions for the Deployment of the Highly Nonlinear MARSIS Antenna Boom Lenticular Joints

The MARSIS antenna booms are constructed using lenticular hinges between straight boom segments in a novel design which allows the booms to be extremely lightweight while retaining a high stiffness and well defined structural properties once they are deployed. Lenticular hinges are elegant in form but are complicated to model as they deploy dynamically and require highly specialized nonlinear techniques founded on carefully measured mechanical properties. Results from component level testing were incorporated into a highly specialized ADAMS model which employed an automated damping algorithm to account for the discontinuous boom lengths formed during the deployment. Additional models with more limited capabilities were also developed in both DADS and ABAQUS to verify the ADAMS model computations and to help better define the numerical behavior of the models at the component and system levels. A careful comparison is made between the ADAMS and DADS models in a series of progressive steps in order to verify their numerical results. Different trade studies considered in the model development are outlined to demonstrate a suitable level of model fidelity. Some model sensitivities to various parameters are explored using subscale and full system models. Finally, some full system DADS models are exercised to illustrate the limitations of traditional modeling techniques for variable geometry systems which were overcome in the ADAMS model.

Mars↗
NASA NTRS1971

Full-scale wind-tunnel tests of a small unpowered jet aircraft with a T-tail

The aerodynamic characteristics of a full scale executive type jet transport aircraft with a T-tail were investigated in a 40 x 80 ft (12.2 by 24.4 meter) wind tunnel (subsonic). Static, longitudinal, and lateral stability, and control characteristics were determined at angles of attack from -2 deg to +42 deg. The aircraft wing had 13 deg of sweep and an aspect ratio of 5.02. The aircraft was tested power off with various wing leading- and trailing-edge high lift devices. The basic configuration was tested with and without such components as engine nacelles, wing tip tanks, and empannage. Hinge-moment data were obtained and downwash angles in the horizontal-tail plane location were calculated. The data were obtained at Reynolds numbers of 4.1 million and 8.7 million based on mean aerodynamic chord. The model had static longitudinal stability through initial stall. Severe tail buffet occurred near the angle of attack for maximum lift. Above initial stall the aircraft had pronounced pitch-up, characteristic of T-tail configurations. A stable trim point was possible at angles of attack between 30 deg and 40 deg (depending on c.g. location and flap setting). Hinge-moment data showed no regions with adverse effects on stick force. Comparisons of wind-tunnel data and flight-test are presented.

Soderman, P. T.↗
NASA NTRS1969

Wagging tail vibration absorber

A 750-foot cantilever length of extendible-tape boom (very low stiffness) was considered as the main system to be damped. A number of tail lengths were tried from 20 feet to 80 feet after which 40 feet was investigated further as a desirable compromise between performance and practical lengths. A 40-foot damping tail produced a damping effect on the main boom for the first mode equivalent in decay rate to 3.1 percent of critical damping. In this case the spring-hinge and tail were tuned to the main boom first mode frequency and the hinge damping was set at 30 percent of critical based on the tail properties. With this same setting, damping of the second mode was .4 percent and the third mode .1 percent.

Barclay, R. G.↗
NASA NTRS1972

Effect of configuration modification on the hypersonic aerodynamic characteristics of a blended delta wing-body entry vehicle

The longitudinal, lateral, and directional aerodynamic characteristics of a delta-wing configuration were obtained experimentally at Mach 20 in helium with Reynolds numbers, based on model length, of 1.5 million and 2.9 million and at a Mach number of 6 in air with a Reynolds number, based on model length, of 4.8 million. The angles of attack varied from 0 deg to 55 deg for two sideslip angles. The effects of the addition of dorsal fins, the removal of wing tip fins, an increase in elevon span, and changes in elevon hinge-line sweep angle are discussed. The unmodified vehicle had a maximum lift-drag ratio of 2.1 at Mach 19 and of 2.4 at Mach 6 with about the same lateral and directional stability level at both Mach numbers. As the Mach number increased from 6 to 20, the longitudinal center of pressure moved forward and more positive elevon deflection was therefore required to maintain a given trim angle. The removal of wing tip fins increased the maximum lift-drag ratio and had a negligible effect on longitudinal stability, but caused directional instability that was not corrected by the dorsal fins examined. The shape of the wing and elevon hinge-line sweep had a large influence on the induced yawing moment due to roll control.

Arrington, J. P.↗
NASA NTRS1972

Analysis of an electrohydraulic aircraft control surface servo and comparison with test results

An analysis of an electrohydraulic aircraft control-surface system is made in which the system is modeled as a lumped, two-mass, spring-coupled system controlled by a servo valve. Both linear and nonlinear models are developed, and the effects of hinge-moment loading are included. Transfer functions of the system and approximate literal factors of the transfer functions for several cases are presented. The damping action of dynamic pressure feedback is analyzed. Comparisons of the model responses with results from tests made on a highly resonant rudder control-surface servo indicate the adequacy of the model. The effects of variations in hinge-moment loading are illustrated.

Edwards, J. W.↗