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At least 505 records · Page 28
All sourcesNASA NTRSDOE OSTI
NASA NTRS1954

A Transonic Wind-Tunnel Investigation of the Longitudinal Aerodynamic Characteristics of a Model of the Lockheed XF-104 Airplane

The transonic longitudinal aerodynamic characteristics of a 0.0858-scale model of the Lockheed XF-104 airplane have been obtained from tests at the Langley 16-foot transonic tunnel. The results of the investigation provide some general information applicable to the transonic properties of thin, low-aspect-ratio, unswept wing configurations utilizing a high horizontal tail . The model employs a horizontal tail mounted at the top of the vertical tail and a wing with an aspect ratio of 2.5, a taper ratio of 0.385, and 3.4-percent-thick airfoil sections. The lift, drag, and static longitudinal pitching moment were measured at Mach numbers from 0.80 t o 1.09 and angles of attack from -2.5 deg to 22.5 deg. Some of the dynamic longitudinal stability properties of the airplane have been predicted from the test results. In addition, some visual flow studies on the wing surfaces obtained at Mach numbers of 0.80 and 1.00 are included. Results of the investigation show that the transonic rise in drag coefficient at zero lift is about 0.030. At high angles of attack, the model becomes longitudinally unstable at Mach numbers from 0.80 t o 0.90, whereas a reduction in static stability is experienced when very high angles of attack are reached at Mach numbers above 0.90. Longitudinal dynamic stability calculations show that the longitudinal control is good at angles of attack below the unstable break in the static pitching-moment curves, but a typical corrective control applied after the occurrence of neutral stability has little effect in averting pitch-up.

Hieser, Gerald↗
NASA NTRS1985

Fuel conservative guidance for shipboard landing of powered-lift STOL aircraft

A computer-simulation study was undertaken to investigate the application of Fuel Conservative Guidance (FCG) techniques, developed at NASA Ames Research Center, to improve the fuel efficiency and minimize recovery time of powered-lift short-takeoff-and-landing (STOL) airplanes operating from aircraft carriers at sea. The FCG system consists of a set of algorithms whose coefficients and parameters limits match those of the Quiet Short-Haul Research Aircraft. When a flightpath is specified by a set of initial conditions for the aircraft and a set of positional waypoints with associated airspeeds, the FCG synthesizes the necessary guidance commands to capture the specified path at any specified waypoint and to optimize fuel consumption and time fo fly along the path. Closed-form expressions are developed for calculating the altitude profile synthesized by the algorithm. Results of this simulation study show that when restrictions on the approach flightpath imposed for manual operation are removed completely, fuel consumption during the approach was reduced by as much as 38 percent (434 lb of fuel) and the time required to fly the flightpath was reduced by as much as 28 percent (209 sec). Savings because of FCG were produced by: (1) shortening the total flight time and distance, and (2) keeping the airspeed high as long as possible to minimize time spent flying in a powered-lift mode.

Warner, D. N., Jr.↗
NASA NTRS1961

Aerodynamic Characteristics over a Mach Number Range of 1.40 to 2.78 of a Rocket-Propelled Airplane Configuration having a Low 52.50 Delta Wing and an Unswept Horizontal Tail

A free-flight investigation of an airplane configuration having a low 52.5 deg. delta wing and an unswept horizontal tail has been conducted over a Mach number range of 1.40 to 2.78. At a fixed tail setting of -3.0 deg., the trim lift coefficient and angle of attack varied from about 0.12 to 0.04 and 3.8 deg. to 2.0 deg., respectively. The base drag was approximately 5 percent of the total drag at trim lift. Lift-curve slope, static longitudinal stability, and damping in pitch were obtained only at Mach numbers of 2.59 t o 2.74. Theoretical calculations of lift-curve slope and aerodynamic-center location were in good agreement with experimental results.

Kehlet, Alan B.↗
NASA NTRS1925

Computation of cantilever airplane wings

The purpose of this treatise is, first of all, the determination of the effect of variously loaded spars on one another, since the neglect of this effect would present an economically very unfavorable computation method. The system of spars and cross-bars alone (whether solid or built-up) does not matter at first, the original assumption being that the spars are rigidly braced by the cross-bars.

WINGS↗
NASA NTRS1972

Aerodynamic parameters of the Navion airplane extracted from flight

An iterative method, which is characterized as a maximum-likelihood minimum-variance technique, was used to extract the aerodynamic parameters of a Navion airplane from flight data. The purposes were to compare the results with parameters obtained from wind-tunnel tests and with results obtained by analog matching the same data, and to develop techniques for application of the parameter extraction program. Results from the study showed that the parameter-extraction program can produce aerodynamic parameters which will permit close estimation of the aircraft time histories used in the extraction process. The program determined an estimate of the standard deviations of the states and parameters. These estimates were used to indicate how well the calculated states fit the flight data and the confidence in the values of the estimated parameters. The study also showed that the values of the parameters were affected by the data and mathematical model used during the extraction process. Because of the lack of confidence in the parameters extracted by use of some of the sets of data, several parameters were estimated by other methods. By using a combination of methods, a set of parameters which gave a fit to the data was obtained.

Suit, W. T.↗
NASA NTRS1930

Experimental Investigation of Aircraft Propellers Exposed to Oblique Air Currents

Two screw propellers, a normal airplane propeller with a pitch/diameter ratio H/D = 0.5 and a helicopter propeller H/D = 0.2 were tested in the large wind tunnel of the Gottigen Aerodynamic Institute. With both propellers the angle between the propeller axis and the direction of the wind was varied from 0 to 90 degrees. In addition to the three force and the three moment components in a wind-fast coordinate system, the corresponding components in a propeller-fast coordinate system, a total of six force and six moment components, together with the propulsive efficiency, were measured or calculated from measurements. The results are given in the form of diagrams and briefly discussed. The propeller models and method of experimentation are described and the symbols are defined.

Flachsbart, O↗
NASA NTRS1948

Supplementary Investigation in the Langley 20-Foot Free-Spinning Tunnel of the Spin and Recovery Characteristics of a 0.057-Scale Model of the Chance Vought XFU-1 Airplane, Ted No. NACA DE 311

An investigation has been conducted in the Langley 20-foot free-spinning tunnel to determine the effects of decreasing the rudder deflection, of decreasing the rudder span, and of differential rudder movements on the spin and recovery characteristics of a 0.057-scale model of the Chance Vought XF7U-1 airplane. The results indicated that decreasing the rudder span or the rudder deflections, individually or jointly, did not seriously alter the spin or recovery characteristics of the model; and recovery by normal use of controls (full rapid rudder reversal followed l/2 to 1 turn later by movement of the stick forward of neutral) remained satisfactory. Linking the original rudders so that the inboard rudder moves from full with the spin to neutral while the outboard rudder moves from neutral to full against the spin will also result in satisfactory spin and recovery characteristics. Calculations of rudder-pedal forces for recovery showed that the expected forces would probably be within the capabilities of a pilot but that it would be advisable to install some type of boost in the control system to insure easy and rapid movement of the rudders.

Berman, Theodore↗
NASA NTRS1976

Flight experience with a digital integrated propulsion control system on an F-111E airplane

A digital integrated propulsion control system (IPCS) installed in the left side of an F-111 E aircraft was tested in flight. The F-111 aircraft was selected for the IPCS program because it incorporated a variable geometry inlet and an afterburning turbofan engine and had two engines, one of which could remain in the normal configuration to ensure flight safety. Flight data were compared with results of tests run in an altitude test chamber. The digital system was found to be capable of duplicating the standard engine and inlet control systems. Instabilities such as inlet buzz and afterburner rumble were detected and controlled. The usefulness of an altitude chamber for developing a software and testing hardware was proven. The flexibility of IPCS was demonstrated when an autothrottle, an in-flight thrust calculation, and a coannular noise study capability were added at the end of the flight tests.

Burcham, F. W., Jr.↗
NASA NTRS1982

Dual wing, swept forward swept rearward wing, and single wing design optimization for high performance business airplanes

An investigation was performed to compare closely coupled dual wing and swept forward swept rearward wing aircraft to corresponding single wing 'baseline' designs to judge the advantages offered by aircraft designed with multiple wing systems. The optimum multiple wing geometry used on the multiple wing designs was determined in an analytic study which investigated the two- and three-dimensional aerodynamic behavior of a wide range of multiple wing configurations in order to find the wing geometry that created the minimum cruise drag. This analysis used a multi-element inviscid vortex panel program coupled to a momentum integral boundary layer analysis program to account for the aerodynamic coupling between the wings and to provide the two-dimensional aerodynamic data, which was then used as input for a three-dimensional vortex lattice program, which calculated the three-dimensional aerodynamic data. The low drag of the multiple wing configurations is due to a combination of two dimensional drag reductions, tailoring the three dimensional drag for the swept forward swept rearward design, and the structural advantages of the two wings that because of the structural connections permitted higher aspect ratios.

Rhodes, M. D.↗
NASA NTRS2000

Airborne Aerosol Closure Studies During PRIDE

The Puerto Rico Dust Experiment (PRIDE) was conducted during June/July of 2000 to study the properties of Saharan dust aerosols transported across the Atlantic Ocean to the Caribbean Islands. During PRIDE, the NASA Ames Research Center six-channel (380 - 1020 nm) airborne autotracking sunphotometer (AATS-6) was operated aboard a Piper Navajo airplane alongside a suite of in situ aerosol instruments. The in situ aerosol instrumentation relevant to this paper included a Forward Scattering Spectrometer Probe (FSSP-100) and a Passive Cavity Aerosol Spectrometer Probe (PCASP), covering the radius range of approx. 0.05 to 10 microns. The simultaneous and collocated measurement of multi-spectral aerosol optical depth and in situ particle size distribution data permits a variety of closure studies. For example, vertical profiles of aerosol optical depth obtained during local aircraft ascents and descents can be differentiated with respect to altitude and compared to extinction profiles calculated using the in situ particle size distribution data (and reasonable estimates of the aerosol index of refraction). Additionally, aerosol extinction (optical depth) spectra can be inverted to retrieve estimates of the particle size distributions, which can be compared directly to the in situ size distributions. In this paper we will report on such closure studies using data from a select number of vertical profiles at Cabras Island, Puerto Rico, including measurements in distinct Saharan Dust Layers. Preliminary results show good agreement to within 30% between mid-visible aerosol extinction derived from the AATS-6 optical depth profiles and extinction profiles forward calculated using 60s-average in situ particle size distributions and standard Saharan dust aerosol refractive indices published in the literature. In agreement with tendencies observed in previous studies, our initial results show an underestimate of aerosol extinction calculated based on the in situ size distributions relative to the extinction obtained from the sunphotometer measurements. However, a more extensive analysis of all available AATS-6 and in situ size distribution data is necessary to ascertain whether the preliminary results regarding the degree of extinction closure is representative of the entire range of dust conditions encountered in PRIDE. Finally, we will compare the spectral extinction measurements obtained in PRIDE to similar data obtained in Saharan dust layers encountered above the Canary Islands during ACE-2 (Aerosol Characterization Experiment) in July 1997. Thus, the evolution of Saharan dust spectral properties during its transport across the Atlantic can be investigated, provided the dust origin and microphysical properties are found to be comparable.

Redemann, Jens↗
NASA NTRS2005

Loads Model Development and Analysis for the F/A-18 Active Aeroelastic Wing Airplane

The Active Aeroelastic Wing airplane was successfully flight-tested in March 2005. During phase 1 of the two-phase program, an onboard excitation system provided independent control surface movements that were used to develop a loads model for the wing structure and wing control surfaces. The resulting loads model, which was used to develop the control laws for phase 2, is described. The loads model was developed from flight data through the use of a multiple linear regression technique. The loads model input consisted of aircraft states and control surface positions, in addition to nonlinear inputs that were calculated from flight-measured parameters. The loads model output for each wing consisted of wing-root bending moment and torque, wing-fold bending moment and torque, inboard and outboard leading-edge flap hinge moment, trailing-edge flap hinge moment, and aileron hinge moment. The development of the Active Aeroelastic Wing loads model is described, and the ability of the model to predict loads during phase 2 research maneuvers is demonstrated. Results show a good match to phase 2 flight data for all loads except inboard and outboard leading-edge flap hinge moments at certain flight conditions. The average load prediction errors for all loads at all flight conditions are 9.1 percent for maximum stick-deflection rolls, 4.4 percent for 5-g windup turns, and 7.7 percent for 4-g rolling pullouts.

Allen, Michael J.↗
NASA NTRS2005

Loads Model Development and Analysis for the F/A-18 Active Aeroelastic Wing Airplane

The Active Aeroelastic Wing airplane was successfully flight-tested in March 2005. During phase 1 of the two-phase program, an onboard excitation system provided independent control surface movements that were used to develop a loads model for the wing structure and wing control surfaces. The resulting loads model, which was used to develop the control laws for phase 2, is described. The loads model was developed from flight data through the use of a multiple linear regression technique. The loads model input consisted of aircraft states and control surface positions, in addition to nonlinear inputs that were calculated from flight-measured parameters. The loads model output for each wing consisted of wing-root bending moment and torque, wing-fold bending moment and torque, inboard and outboard leading-edge flap hinge moment, trailing-edge flap hinge moment, and aileron hinge moment. The development of the Active Aeroelastic Wing loads model is described, and the ability of the model to predict loads during phase 2 research maneuvers is demonstrated. Results show a good match to phase 2 flight data for all loads except inboard and outboard leading-edge flap hinge moments at certain flight conditions. The average load prediction errors for all loads at all flight conditions are 9.1 percent for maximum stick-deflection rolls, 4.4 percent for 5-g windup turns, and 7.7 percent for 4-g rolling pullouts.

Allen, Michael J.↗
NASA NTRS1977

A theoretical analysis of airplane longitudinal stability and control as affected by wind shear

The longitudinal equations of motion with wind shear terms were used to analyze the stability and motions of a jet transport. A positive wind shear gives a decreasing head wind or changes a head wind into a tail wind. A negative wind shear gives a decreasing tail wind or changes a tail wind into a head wind. It was found that wind shear had very little effect on the short period mode and that negative wind shear, although it affected the phugoid, did not cause stability problems. On the other hand, it was found that positive wind shear can cause the phugoid to become aperiodic and unstable. In this case, a stability boundary for the phugoid was found that is valid for most aircraft at all flight speeds. Calculations of aircraft motions confirmed the results of the stability analysis. It was found that a flight path control automatic pilot and an airspeed control system provide good control in all types of wind shear. Appendixes give equations of motion that include the effects of downdrafts and updrafts and extend the longitudinal equations of motion for shear to six degrees of freedom.

Sherman, W. L.↗
NASA NTRS1977

Flight evaluation of an advanced technology light twin-engine airplane (ATLIT)

Project organization and execution, airplane description and performance predictions, and the results of the flight evaluation of an advanced technology light twin engine airplane (ATLIT) are presented. The ATLIT is a Piper PA-34-200 Seneca I modified by the installation of new wings incorporating the GA(W)-1 (Whitcomb) airfoil, reduced wing area, roll control spoilers, and full span Fowler flaps. The conclusions for the ATLIT evaluation are based on complete stall and roll flight test results and partial performance test results. The Stalling and rolling characteristics met design expectations. Climb performance was penalized by extensive flow separation in the region of the wing body juncture. Cruise performance was found to be penalized by a large value of zero lift drag. Calculations showed that, with proper attention to construction details, the improvements in span efficiency and zero lift drag would permit the realization of the predicted increases in cruising and maximum rate of climb performance.

Holmes, B. J.↗
NASA NTRS1921

Performance of a Liberty 12 airplane engine

In cooperation with the Engineering Division of the Air Service of the United States Army, a Liberty-12 engine has been tested at the Bureau of Standards. The program of tests was planned to yield that information considered most important in determining the value of the engine for aviation. Full power runs were made at the ground, at 25,000 feet, and at several intermediate altitudes. To determine the mechanical efficiency of the engine, friction horsepower was measured at the ground and at 15,000 feet. As a basis for predicting engine performance with a propeller, a series of tests was made in which the dynamometer load and engine throttle were adjusted at each speed to simulate the engine load which would be imposed at that speed by a propeller operating under normal full load at 1,700 r.p.m. Among the quantities calculated from the test measurements are: brake horsepower; break mean effective pressure; fuel consumption; mixture ratio; mechanical, thermal, and volumetric efficiency; and the percentage of the heat in the fuel appearing in the jacket water and in the exhaust. Jacket water temperature, oil temperature, manifold pressure, etc., are recorded to show the conditions under which the test was made.

Sparrow, S W↗
NASA NTRS2007

Computing a Stability Spectrum by Use of the HHT

The Hilbert-Huang transform (HHT) is part of the mathematical basis of a method of calculating a stability spectrum. This method can be regarded as an extended and improved version of a prior HHT-based method of calculating a damping spectrum. In the prior method, information on positive damping (which leads to stability) and negative damping (which leads to instability) becomes mixed into a single squared damping loss factor. Hence, there is no way to distinguish between stability and instability by examining a damping spectrum. In contrast, in the present stabilityspectrum method, information on the mathematical sign of the damping is retained, making it possible to identify reThe Hilbert-Huang transform (HHT) is part of the mathematical basis of a method of calculating a stability spectrum. This method can be regarded as an extended and improved version of a prior HHT-based method of calculating a damping spectrum. In the prior method, information on positive damping (which leads to stability) and negative damping (which leads to instability) becomes mixed into a single squared damping loss factor. Hence, there is no way to distinguish between stability and instability by examining a damping spectrum. In contrast, in the present stabilityspectrum method, information on the mathematical sign of the damping is retained, making it possible to identify regions of instability in a spectrum. This method is expected to be especially useful for analyzing vibration-test data for the purpose of predicting vibrational instabilities in structures (e.g., flutter in airplane wings).

Huang, Norden E.↗
NASA NTRS1985

A fundamental study of the sticking of insect residues to aircraft wings

The aircraft industry has long been concerned with the increase of drag on airplanes due to fouling of the wings by insects. The present research studied the effects of surface energy and surface roughness on the phenomenon of insect sticking. Aluminum plates of different roughnesses were coated with thin films of polymers with varying surface energies. The coated plates were attached to a custom jig and mounted on top of an automobile for insect collection. Contact angle measurements, X-ray photoelectron spectroscopy and specular reflectance infrared spectroscopy were used to characterize the surface before and after the insect impact experiments. Scanning electron microscopy showed the topography of insect residues on the exposed plates. Moments were calculated in order to find a correlation between the parameters studied and the amount of bugs collected on the plates. An effect of surface energy on the sticking of insect residues was demonstrated.

Eiss, N. S., Jr.↗
NASA NTRS1987

Wind-tunnel investigation of a general aviation airplane equipped with a high aspect-ratio, natural laminar-flow wing

A wind tunnel investigation has been conducted to evaluate the aerodynamic performance, stability, and control characteristics of a full-scale general aviation aircraft equipped with an advanced natural laminar flow wing. Force and moment data were measured over wide angle-of-attack and sideslip ranges and at Reynolds numbers from 1.4 million to 2.4 million based on the mean aerodynamic chord. Additional measurements determined the condition of the wing boundary layer and the wing stalling characteristics. The results showed that large regions of natural laminar flow existed on the wing which would significantly enhance the cruise performance of the configuration. Artificially tripping the boundary layer did not significantly affect the lift, stability, and control characteristics. Leading-edge modifications were effective in improving the stall resistance of the configuration, and resulted in only minor penalties in calculated cruise performance.

Murri, Daniel G.↗