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At least 1,567 records · Page 87

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NASA NTRS1979

The Influence of Sweep on the Aerodynamic Loading of an Oscillating NACA0012 Airfoil

The effect of sweep on the dynamic response of the NACA 0012 airfoil was investigated. Unsteady chordwise distributed pressure data were obtained from a tunnel spanning wing equipped with 21 single surface transducers (13 on the suction side and 8 on the pressure side of the airfoil). The pressure data were obtained at pitching amplitudes of 8 and 10 degrees over a tunnel Mach number range of 0.10 to 0.46 and a pitching frequency range of 2.5 to 10.6 cycles per second. The wing was oscillated in the unswept and swept positions about the quarter-chord pivot axis relative to mean incidence angle settings of 0, 9, 12, and 15 degrees. A compilation of all the response data obtained during the test program is presented. These data are in the form of normal force, chord force, lift force, pressure drag, and moment hysteresis loops derived from chordwise integrations of the unsteady pressure distributions. The hysteresis loops are organized in two main sections. In the first section, the loop data are arranged to show the effect of sweep (lambda = 0 and 30 deg) for all available combinations of mean incidence angle, pitching amplitude, reduced frequency, and chordwise Mach number. The second section shows the effect of chordwise Mach number (MC = 0.30 and MC = 0.40) on the swept wing response for all available combinations of mean incidence angle, pitching amplitude, and reduced frequency.

St.hilaire, A. O.↗

NASA NTRS1927

Researches on ailerons and especially on the test loads to which they should be subjected

Aileron calculations have hitherto given greatly differing results according to different authors. It seems to be the general opinion that it is only necessary to give the ailerons such dimensions that the airplane can maneuver well, that the stresses they must undergo are relatively small, and that they are strong enough if their framework is of the order of strength as the wings to which they are attached. This article will show that the problem is really quite complex and that it should receive more attention.

AILERONS - TEST LOADS↗

NASA NTRS1977

Optimal control alleviation of tilting proprotor gust response

Optimal control theory is applied to the design of a control system for alleviation of the gust response of tilting proprotor aircraft. Using a proprotor and cantilever wing analytical model, the uncontrolled and controlled gust response is examined over the entire operating range of the aircraft except for hover: helicopter mode, conversion, and airplane mode flight. Substantial improvements in the loads, ride quality, and aeroelastic stability are possible with a properly designed controller. A single controller, nominally optimal only at the design point speed (160 knots here), operated efficiently over the entire speed range, with the possible exception of very low speed in helicopter mode. Kalman-Bucy filters were used as compensation networks to provide state estimates from various measurements in the system. Efficient control requires the measurement of the wing motion, rotor speed perturbation, and tip-path-plane tilt.

Johnson, W.↗

NASA NTRS2012

Aeroelastic Modeling of Elastically Shaped Aircraft Concept via Wing Shaping Control for Drag Reduction

Lightweight aircraft design has received a considerable attention in recent years as a means for improving cruise efficiency. Reducing aircraft weight results in lower lift requirements which directly translate into lower drag, hence reduced engine thrust requirements during cruise. The use of lightweight materials such as advanced composite materials has been adopted by airframe manufacturers in current and future aircraft. Modern lightweight materials can provide less structural rigidity while maintaining load-carrying capacity. As structural 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. Abstract This paper describes a recent aeroelastic modeling effort for an elastically shaped aircraft concept (ESAC). The aircraft model is based on the rigid-body generic transport model (GTM) originally developed at NASA Langley Research Center. The ESAC distinguishes itself from the GTM in that it is equipped with highly flexible wing structures as a weight reduction design feature. More significantly, the wings are outfitted with a novel control effector concept called variable camber continuous trailing edge (VCCTE) flap system for active control of wing aeroelastic deflections to optimize the local angle of attack of wing sections for improved aerodynamic efficiency through cruise drag reduction and lift enhancement during take-off and landing. The VCCTE flap is a multi-functional and aerodynamically efficient device capable of achieving high lift-to-drag ratios. The flap system is comprised of three chordwise segments that form the variable camber feature of the flap and multiple spanwise segments that form a piecewise continuous trailing edge. By configuring the flap camber and trailing edge shape, drag reduction could be achieved. Moreover, some parts of the flap system can be made to have a high frequency response for roll control, gust load alleviation, and aeroservoelastic (ASE) modal suppression control. Abstract The aeroelastic model of the ESAC is based on one-dimensional structural dynamic theory that captures the aeroelastic deformation of a wing structure in a combined motion that involves flapwise bending, chordwise bending, and torsion. The model includes the effect of aircraft propulsion due to wing flexibility which causes the propulsive forces and moments to couple with the wing elastic motion. Engine mass is also accounted in the model. A fuel management model is developed to describe the wing mass change due to fuel usage in the main tank and wing tanks during cruise. Abstract The model computes both static and dynamic responses of the wing structures. The static aeroelastic deflections are used to estimate the effect of wing flexibility on induced drag and the potential drag reduction by the VCCTE flap system. A flutter analysis is conducted to estimate the flutter speed boundary. Gust load alleviation via adaptive control has been recently investigated to address flexibility of aircraft structures. A multi-objective flight control approach is presented for drag reduction control. The approach is based on an optimal control framework using a multi-objective cost function. Future studies will demonstrate the potential benefits of the approach.

Aeroelasticity↗

NASA NTRS1991

Optimization of composite sandwich cover panels subjected to compressive loadings

An analysis and design method is presented for the design of composite sandwich cover panels that includes transverse shear effects and damage tolerance considerations. This method is incorporated into an optimization program called SANDOP (SANDwich OPtimization). SANDOP is used in the present study to design optimized composite sandwich cover panels for transport aircraft wing applications as a demonstration of its capabilities. The results of this design study indicate that optimized composite sandwich cover panels have approximately the same structural efficiency as stiffened composite cover panels designed to identical constraints. Results indicate that inplane stiffness requirements have a large effect on the weight of these composite sandwich cover panels at higher load levels. Increasing the maximum allowable strain and the upper percentage limit of the 0 degree and plus or minus 45 degree plies can yield significant weight savings. The results show that the structural efficiency of these optimized composite sandwich cover panels is relatively insensitive to changes in core density.

Cruz, Juan R.↗

NASA NTRS2011

Distributed Aerodynamic Sensing and Processing Toolbox

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.

Brenner, Martin↗

NASA NTRS1973

Pressures and temperatures on the lower surfaces of an externally blown flap system during full-scale ground tests

Full-scale ground tests of an externally blown flap system were made using the wing of an F-111B airplane and a CF700 engine. Pressure and temperature distributions were determined on the undersurface of the wing, vane, and flap for two engine exhaust nozzles (conical and daisy) at several engine power levels and engine/wing positions. The test were made with no airflow over the wing. The wing sweep angle was fixed at e6 deg; and the angle of incidence between the engine and the wing was fixed at 3 deg; and the flap was in the retracted, deflected 35 deg, and deflected 60 deg positions. The pressure load obtained by integrating the local pressures on the undersurface of the flap, F sub p was approximately three times greater at the 60 deg flap position than at the 35 deg flap position. At the 60 deg flap position, F sub p was between 40 percent and 55 percent of the engine thrust over the measured range of thrust. More than 90 percent of F sub p was contained within plus or minus 20 percent of the flap span centered around the engine exhaust centerline with both nozzle configurations. Maximum temperatures recorded on the flaps were 218 C (424 F) and 180 C (356 F) for the conical and daisy nozzles, repectively,

Hughes, D. L.↗

NASA NTRS1981

Vortex-flow aerodynamics - An emerging design capability

Promising current theoretical and simulational developments in the field of leading edge vortex-generating delta, arrow ogival wings are reported, along with the history of theory and experiment leading to them. The effects of wing slenderness, leading edge nose radius, Mach number and incidence variations, and planform on the onset of vortex generation and redistribution of aerodynamic loads are considered. The range of design possibilities in this field are consequential for the future development of strategic aircraft, supersonic transports and commercial cargo aircraft which will possess low-speed, high-lift capability by virtue of leading edge vortex generation and control without recourse to heavy and expensive leading edge high-lift devices and compound airfoils. Attention is given to interactive graphics simulation devices recently developed.

Campbell, J. F.↗

NASA NTRS1980

Integrated application of active controls (IAAC) technology to an advanced subsonic transport project. Conventional baseline configuration study

Characteristics of the U.S. domestic fleet were evaluated to determine the mission characteristics that would have the most impact on U. S. transport fuel use in the future. This resulted in selection of a 197-passenger (plus cargo), about 3710-km (2000 nmi) mission. The existing data base was reviewed and additional analysis was conducted as necessary to complete the technical descriptions. The resulting baseline configuration utilizes a double-lobe, but nearly circular, body with seven-abreast seating. External characteristics feature an 8.71 aspect ratio, 31.5-degree sweep wing, a T-tail empennage, and a dual CF6-6D2, wing-mounted engine arrangement. It provides for 22 LD-2 or 11 LD-3 containers plus bulk cargo in the lower lobe. Passenger/cargo loading, servicing provisions, taxi/takeoff speeds, and field length characteristics are all compatible with accepted airline operations and regulatory provisions. The baseline configuration construction uses conventional aluminum structure except for advanced aluminum alloys and a limited amount of graphite epoxy secondary structure. Modern systems are used, including advanced guidance, navigation, and controls which emphasize application of digital electronics and advanced displays.

Source record↗

NASA NTRS1984

Unconventional missile concepts from consideration of varied mission requirements

The characteristics of a series of unconventional missile configurations are reviewed in the light of specific mission requirements, taking volumetric efficiency and aerodynamic performance into account. Cones with and without delta wings, thick delta wings, a semiconical body with delta wings, a ring wing, a flat body with half-ring or swept-parasol wings, a parasol-wing/body with a high/low wing, and monoplanar missiles with circular or elliptical bodies are described and illustrated. Mission requirements discussed include tactical and strategic penetration, maneuverability, load-carrying capability, low detectability, low cost, and ease of carriage and storage. In general, delta configurations offer high speed and good load-carrying ability at low altitudes; parasol configurations give high lift and low drag at high altitude and speed; and monoplanar-elliptical configurations have high lift and low drag with good maneuvering and stability.

Spearman, M. L.↗

NASA NTRS1987

Static aeroelastic effects on the flutter of a supercritical wing

It is well known that wings with supercritical airfoils generally have lower transonic flutter speeds than similar wings with conventional airfoils and that small increases in angle of attack from zero and the accompanying static aeroelastic deformations have further detrimental effects on transonic flutter. The results of an effort to calculate the effects of angle of attack and the associated aeroelastic deformation on the flutter of a highly swept supercritical wing (TF 8A) by use of the modified strip analysis employed in previous studies of this wing are presented. The spanwise distributions of steady-state section lift-curve slope and aerodynamic center required as input for these calculations were obtained from static aeroelastic calculations for the wing by use of the FL022 transonic code and an assumed dynamic pressure as that used to calculate the statically deformed shape and loading about which the flutter oscillation occurs. The results show that the unconventional backward turn of the transonic dip in the experimental flutter boundary for angles of attack greater than zero is caused by variations in mass ratio and not by static aeroelastic deformation, although inclusion of the latter appears to be required for quantitative accuracy in the calculations.

Yates, E. Carson, Jr.↗

NASA NTRS1986

Flow rate and trajectory of water spray produced by an aircraft tire

One of the risks associated with wet runway aircraft operation is the ingestion of water spray produced by an aircraft's tires into its engines. This problem can be especially dangerous at or near rotation speed on the takeoff roll. An experimental investigation was conducted in the NASA Langley Research Center Hydrodynamics Research Facility to measure the flow rate and trajectory of water spray produced by an aircraft nose tire operating on a flooded runway. The effects of various parameters on the spray patterns including distance aft of nosewheel, speed, load, and water depth were evaluated. Variations in the spray pattern caused by the airflow about primary structure such as the fuselage and wing are discussed. A discussion of events in and near the tire footprint concerning spray generation is included.

Daugherty, Robert H.↗

NASA NTRS1986

Buckling characteristics of hypersonic aircraft wing tubular panels

The buckling characteristics of Rene 41 tubular panels installed as wing panels on a hypersonic wing test structure (HWTS) were determined nondestructively through use of a force/stiffness technique. The nondestructive buckling tests were carried out under different combined load conditions and different temperature environments. Two panels were subsequently tested to buckling failure in a universal tension compression testing machine. In spite of some data scattering because of large extrapolations of data points resulting from termination of the test at a somewhat low applied load, the overall test data correlated fairly well with theoretically predicted buckling interaction curves. The structural efficiency of the tubular panels was slightly higher than that of the beaded panels which they replaced.

Ko, William L.↗

NASA NTRS1939

Downwash and Wake Behind Plain and Flapped Airfoils

Extensive experimental measurements have been made of the downwash angles and the wake characteristics behind airfoils with and without flaps and the data have been analyzed and correlated with the theory. A detailed study was made of the errors involved in applying lifting-line theory, such as the effects of a finite wing chord, the rolling-up of the trailing vortex sheet, and the wake. The downwash angles, as computed from the theoretical span load distribution by means of the Biot-Savart equation, were found to be in satisfactory agreement with the experimental results. The rolling-up of the trailing vortex sheet may be neglected, but the vertical displacement of the vortex sheet requires consideration. By the use of a theoretical treatment indicated by Prandtl, it has been possible to generalize the available experimental results so the predictions can be made of the important wake parameters in terms of the distance behind the airfoil trailing edge and the profile-drag coefficient. The method of application of the theory to design and the satisfactory agreement between predicted and experimental results when applied to an airplane are demonstrated.

Silverstein, Abe↗

NASA NTRS1999

Rotationally Adaptive Flight Test Surface

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.

Barrett, Ron↗

NASA NTRS2001

Full-Span Tiltrotor Aeroacoustic Model (TRAM) Overview and 40- by 80-Foot Wind Tunnel Test

Most helicopter data trends cannot be extrapolated to tiltrotors because blade geometry and aerodynamic behavior, as well as rotor and fuselage interactions, are significantly different for tiltrotors. A tiltrotor model has been developed to investigate the aeromechanics of tiltrotors, to develop a comprehensive database for validating tiltrotor analyses, and to provide a research platform for supporting future tiltrotor designs. The Full-Span Tiltrotor Aeroacoustic Model (FS TRAM) is a dual-rotor, powered aircraft model with extensive instrumentation for measurement of structural and aerodynamic loads. This paper will present the Full-Span TRAM test capabilities and the first set of data obtained during a 40- by 80-Foot Wind Tunnel test conducted in late 2000 at NASA Ames Research Center. The Full-Span TRAM is a quarter-scale representation of the V-22 Osprey aircraft, and a heavily instrumented NASA and U.S. Army wind tunnel test stand. Rotor structural loads are monitored and recorded for safety-of-flight and for information on blade loads and dynamics. Left and right rotor balance and fuselage balance loads are monitored for safety-of-flight and for measurement of vehicle and rotor aerodynamic performance. Static pressure taps on the left wing are used to determine rotor/wing interactional effects and rotor blade dynamic pressures measure blade airloads. All of these measurement capabilities make the FS TRAM test stand a unique and valuable asset for validation of computational codes and to aid in future tiltrotor designs. The Full-Span TRAM was tested in the NASA Ames Research Center 40- by 80-Foot Wind Tunnel from October through December 2000. Rotor and vehicle performance measurements were acquired in addition to wing pressures, rotor acoustics, and Laser Light Sheet (LLS) flow visualization data. Hover, forward flight, and airframe (rotors off) aerodynamic runs were performed. Helicopter-mode data were acquired during angle of attack and thrust sweeps for a variety of tunnel speeds. Wake geometry images were acquired using LLS photographs and suggest dual tip vortex formation at low thrust conditions. The full paper will include comparisons to isolated-rotor TRAM data acquired at the Duits-Nederlandse Windtunnel (DNW) in 1998. The FS TRAM has been established as a valuable national asset for tiltrotor research. Data reduction and analysis of the 40- by 80-Foot Wind Tunnel test results are underway. Follow-on testing of the FS TRAM is currently being planned for the NASA Ames 80- by 120-Foot Wind Tunnel in late 2001.

McCluer, Megan S.↗

NASA NTRS1961

Fatigue Investigation of Full-Scale Wing Panels of 7075 Aluminum Alloy

Results are presented of a fatigue investigation conducted using 18 outer wing panels of T-29A airplanes. Constant-amplitude tests were performed using the resonant-frequency method at three different alternating load levels superposed on a 1 g, or level-flight mean load. Information is presented concerning crack location, cycles to crack initiation, crack propagation, and residual static strength.

Castle, Claude B.↗

NASA NTRS2005

STS-114: Discovery Post MMT Briefing

Wayne Hale, Space Shuttle Deputy Program Manager and Steve Poulos, Manager Orbit Project Office at Johnson Space Center is seen during a post Mission Management Team (MMT) briefing. The purpose of this briefing is to talk about the status of the Space Shuttle Discovery Orbiter, the Thermal Protection System (TPS) and the External Tank. Hale presents pictures of the missing foam off of the Hydrogen Protuberance Air Load (PAL) ramp and trajectory debris particles. Poulos explains in detail diagrams of the reinforced carbon-carbon system, possible scuff on the wing panel, coating damage, protruding gap filler and damaged tile blanket. Poulos tells what concerns him the most in terms of tile damage to the Space Shuttle Discovery.

Source record↗