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At least 487 records · Page 27

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NASA NTRS1956

Drag Interference Between a Pointed Cylindrical Body and Triangular Wings of Various Aspect Ratios at Mach Numbers of 1.50 and 2.02

The drag of a body alone, six triangular wings of various aspect ratios, and the combinations were measured at Mach numbers of 1.50 and 2.02 at a Reynolds number of 5.5 million (based on the body length). The experimental drag-interference results were in accordance with calculations based on NACA RM A9E19, 1949, with skin-friction effects taken into account, the interference effect being principally the result of fixing transition on the body by adding a wing.

BODIES↗

NASA NTRS1999

Noise Certification Predictions for FJX-2-Powered Aircraft Using Analytic Methods

Williams International Co. is currently developing the 700-pound thrust class FJX-2 turbofan engine for the general Aviation Propulsion Program's Turbine Engine Element. As part of the 1996 NASA-Williams cooperative working agreement, NASA agreed to analytically calculate the noise certification levels of the FJX-2-powered V-Jet II test bed aircraft. Although the V-Jet II is a demonstration aircraft that is unlikely to be produced and certified, the noise results presented here may be considered to be representative of the noise levels of small, general aviation jet aircraft that the FJX-2 would power. A single engine variant of the V-Jet II, the V-Jet I concept airplane, is also considered. Reported in this paper are the analytically predicted FJX-2/V-Jet noise levels appropriate for Federal Aviation Regulation certification. Also reported are FJX-2/V-Jet noise levels using noise metrics appropriate for the propeller-driven aircraft that will be its major market competition, as well as a sensitivity analysis of the certification noise levels to major system uncertainties.

Berton, Jeffrey J.↗

NASA NTRS1961

Skid Landings of Airplanes on Rocker-Type Fuselages

A study is made of the landing of an airplane on a fuselage with "planned" curvature of its lower surface. Initial contact is considered to stop the vertical motion of a point remote from the center of gravity, thus causing rocking on the curved lower surface which converts sinking-speed energy into angular energy in pitch for dissipation by damping forces. Analysis is made of loads and motions for a given fuselage shape, and the contours required to give desired load histories are determined. Most of the calculations involve initial contact at the tail, but there are two cases of unflared landings with initial contact at the nose. The calculations are checked experimentally for the tail - low case.

Mayo, Wilbur L.↗

NASA NTRS1981

Qualitative comparison of calculated turbulence responses with wind-tunnel measurements for a DC-10 derivative wing with an active control system

This paper presents comparisons of analytically-predicted and experimental turbulence responses of a wind-tunnel model of a DC-10 derivative wing equipped with an active control system. The active control system was designed for the purpose of flutter suppression, but it had the additional benefit of alleviating gust loads (wing bending moment) by about 25 percent. Comparisons of various wing responses are presented for variations in active-control-system parameters and tunnel speed. The analytical turbulence responses were obtained using DYLOFLEX, a computer program for dynamic loads analyses of flexible airplanes with active controls. In general, the analytical predictions agreed reasonably well with the experimental data.

Perry, B., III↗

NASA NTRS1978

Boundary layer, skin friction, and boattail pressure measurements from the YF-12 airplane at Mach numbers up to 3

In-flight measurements of boundary layer and skin friction data were made on YF-12 airplanes for Mach numbers between 2.0 and 3.0. Boattail pressures were also obtained for Mach numbers between 0.7 and 3.0 with Reynolds numbers up to four hundred million. Boundary layer data measured along the lower fuselage centerline indicate local displacement and momentum thicknesses can be much larger than predicted. Skin friction coefficients measured at two of five lower fuselage stations were significantly less than predicted by flat plate theory. The presence of large differences between measured boattail pressure drag and values calculated by a potential flow solution indicates the presence of vortex effects on the upper boattail surface. At both subsonic and supersonic speeds, pressure drag on the longer of two boattail configurations was equal to or less than the pressure drag on the shorter configuration. At subsonic and transonic speeds, the difference in the drag coefficient was on the order of 0.0008 to 0.0010. In the supersonic cruise range, the difference in the drag coefficient was on the order of 0.002. Boattail drag coefficients are based on wing reference area.

Fisher, D. F.↗

NASA NTRS1992

Simulation model of a twin-tail, high performance airplane

The mathematical model and associated computer program to simulate a twin-tailed high performance fighter airplane (McDonnell Douglas F/A-18) are described. The simulation program is written in the Advanced Continuous Simulation Language. The simulation math model includes the nonlinear six degree-of-freedom rigid-body equations, an engine model, sensors, and first order actuators with rate and position limiting. A simplified form of the F/A-18 digital control laws (version 8.3.3) are implemented. The simulated control law includes only inner loop augmentation in the up and away flight mode. The aerodynamic forces and moments are calculated from a wind-tunnel-derived database using table look-ups with linear interpolation. The aerodynamic database has an angle-of-attack range of -10 to +90 and a sideslip range of -20 to +20 degrees. The effects of elastic deformation are incorporated in a quasi-static-elastic manner. Elastic degrees of freedom are not actively simulated. In the engine model, the throttle-commanded steady-state thrust level and the dynamic response characteristics of the engine are based on airflow rate as determined from a table look-up. Afterburner dynamics are switched in at a threshold based on the engine airflow and commanded thrust.

Buttrill, Carey S.↗

NASA NTRS1982

Use of optimization to predict the effect of selected parameters on commuter aircraft performance

The relationships between field length and cruise speed and aircraft direct operating cost were determined. A gradient optimizing computer program was developed to minimize direct operating cost (DOC) as a function of airplane geometry. In this way, the best airplane operating under one set of constraints can be compared with the best operating under another. A constant 30-passenger fuselage and rubberized engines based on the General Electric CT-7 were used as a baseline. All aircraft had to have a 600 nautical mile maximum range and were designed to FAR part 25 structural integrity and climb gradient regulations. Direct operating cost was minimized for a typical design mission of 150 nautical miles. For purposes of C sub L sub max calculation, all aircraft had double-slotted flaps but with no Fowler action.

Wells, V. L.↗

NASA NTRS1984

Propagation of propeller tone noise through a fuselage boundary layer

In earlier experimental and analytical studies, it was found that the boundary layer on an aircraft could provide significant shielding from propeller noise at typical transport airplane cruise Mach numbers. In this paper a new three-dimensional theory is described that treats the combined effects of refraction and scattering by the fuselage and boundary layer. The complete wave field is solved by matching analytical expressions for the incident and scattered waves in the outer flow to a numerical solution in the boundary layer flow. The model for the incident waves is a near-field frequency-domain propeller source theory developed previously for free field studies. Calculations for an advanced turboprop (Prop-Fan) model flight test at 0.8 Mach number show a much smaller than expected pressure amplification at the noise directivity peak, strong boundary layer shielding in the forward quadrant, and shadowing around the fuselage. Results are presented showing the difference between fuselage surface and free-space noise predictions as a function of frequency and Mach number. Comparison of calculated and measured effects obtained in a Prop-Fan model flight test show good agreement, particularly near and aft of the plane of rotation at high cruise Mach number.

Hanson, D. B.↗

NASA NTRS1974

Determining spacecraft and airplane angular orientation from star photographs

The advantages of determining spacecraft angular orientation from star photographs include the documental nature and objectivity of the photographic image and also the high accuracy of construction of the reference system based on the stars. Use of the method presupposes installation aboard the vehicle of a specialized photographic camera adapted for photographing the stellar sky. From the mathematical viewpoint, the essence of the method amounts to: (1) Certain directions in space are identified and their direction cosines are determined in a specified coordinate system. (2) The direction cosines of these same directions are determined relative to a coordinate system fixed with the vehicle. (3) The sought angular orientation of the vehicle is calculated using the coupling equations between the specified and vehicle-fixed coordinate systems. In principle, vehicle angular orientation can be determined relative to any given reference system.

Elbakyan, K. I.↗

NASA NTRS1989

Development and application of a program to calculate transonic flow around an oscillating three-dimensional wing using finite difference procedures

A finite difference method for solving the unsteady transonic flow about harmonically oscillating wings is investigated. The procedure is based on separating the velocity potential into steady and unsteady parts and linearizing the resulting unsteady differential equation for small disturbances. The differential equation for the unsteady potential is linear with spatially varying coefficients and with the time variable eliminated by assuming harmonic motion. Difference equations are derived for harmonic transonic flow to include a coordinate transformation for swept and tapered planforms. A pilot program is developed for three-dimensional planar lifting surface configurations (including thickness) for the CRAY-XMP at Boeing Commercial Airplanes and for the CYBER VPS-32 at the NASA Langley Research Center. An investigation is made of the effect of the location of the outer boundaries on accuracy for very small reduced frequencies. Finally, the pilot program is applied to the flutter analysis of a rectangular wing.

Weatherill, Warren H.↗

NASA NTRS1959

Analysis and Evaluation of Supersonic Underwing Heat Addition

The linearized theory for heat addition under a wing has been developed to optimize wing geometry, heat addition, and angle of attack. The optimum wing has all of the thickness on the underside of the airfoil, with maximum-thickness point well downstream, has a moderate thickness ratio, and operates at an optimum angle of attack. The heat addition is confined between the fore Mach waves from under the trailing surface of the wing. By linearized theory, a wing at optimum angle of attack may have a range efficiency about twice that of a wing at zero angle of attack. More rigorous calculations using the method of characteristics for particular flow models were made for heating under a flat-plate wing and for several wings with thickness, both with heat additions concentrated near the wing. The more rigorous calculations yield in practical cases efficiencies about half those estimated by linear theory. An analysis indicates that distributing the heat addition between the fore waves from the undertrailing portion of the wing is a way of improving the performance, and further calculations appear desirable. A comparison of the conventional ramjet-plus wing with underwing heat addition when the heat addition is concentrated near the wing shows the ramjet to be superior on a range basis up to Mach number of about B. The heat distribution under the wing and the assumed ramjet and airframe performance may have a marked effect on this conclusion. Underwing heat addition can be useful in providing high-altitude maneuver capability at high flight Mach numbers for an airplane powered by conventional ramjets during cruise.

wings↗

NASA NTRS1973

Flight test measurements and analysis of sonic boom phenomena near the shock wave extremity

The sonic boom flight test program conducted at Jackass Flats, Nevada, during the summer and fall of 1970 consisted of 121 sonic-boom-generating flights over the 1500 ft instrumented BREN tower. This test program was designed to provide information on several aspects of sonic boom, including caustics produced by longitudinal accelerations, caustics produced by steady flight near the threshold Mach number, sonic boom characteristics near lateral cutoff, and the vertical extent of shock waves attached to near-sonic airplanes. The measured test data, except for the near-sonic flight data, were analyzed in detail to determine sonic boom characteristics for these flight conditions and to determine the accuracy and the range of validity of linear sonic boom theory. The caustic phenomena observed during the threshold Mach number flights and during the transonic acceleration flights are documented and analyzed in detail. The theory of geometric acoustics is shown to be capable of predicting shock wave-ground intersections, and current methods for calculating sonic boom pressure signature away from caustics are shown to be reasonably accurate.

Haglund, G. T.↗

NASA NTRS1982

Assessment of advanced technologies for high performance single-engine business airplanes

The prospects for significantly increasing the fuel efficiency and mission capability of single engine business aircraft through the incorporation of advanced propulsion, aerodynamics and materials technologies are explored. It is found that turbine engines cannot match the fuel economy of the heavier rotary, diesel and advanced spark reciprocating engines. The rotary engine yields the lightest and smallest aircraft for a given mission requirement, and also offers greater simplicity and a multifuel capability. Great promise is also seen in the use of composite material primary structures in conjunction with laminar flow wing surfaces, a pusher propeller and conventional wing-tail configuration. This study was conducted with the General Aviation Synthesis Program, which can furnish the most accurate mission performance calculations yet obtained.

Kohlman, D. L.↗

NASA NTRS1975

Two-stage, low noise advanced technology fan. 5: Acoustic final report

The NASA Q2S(quiet two-stage) fan is a 0.836m (32.9 in.) diameter model of the STF 433 engine fan, selected in a 1972 study for an Advanced Technology Transport (ATT) airplane. Noise-control features include: low tip speed, moderate stage pressure rise, large blade-vane spacings, no inlet guide vanes, and optimum blade and vane numbers. Tests were run on the baseline Q2S fan with standard inlet and discharge ducts. Further tests were made of a translating centerbody sonic inlet device and treated discharge ducts. Results were scaled to JT8D and JT3D engine fan size for comparison with current two-stage fans, and were also scaled to STF 433 fan size to compare calculated ATT flyover noise with FAR 36 limits. Baseline Q2S results scaled to JT8D and JT3D engine fan sizes showed substantial noise reductions. Calculated unsuppressed baseline ATT flyovers averaged about 2.5 EPNdB below FAR 36 limits. Using measured sonic inlet results, scaled baseline Q2S fan results, and calculated attenuations for a 1975 technology duct liner, projected flyover noise calculations for the ATT averaged about FAR 36 limits minus 10 EPNdB. Advances in suppression technology required to meet the 1985 goal of FAR 36 limits minus 20 EPNdB are discussed.

Sofrin, T. G.↗

NASA NTRS1979

Airborne spacecraft - A remotely powered, high-altitude RPV for environmental applications

A high-altitude, unmanned, propeller-driven electric airplane is proposed for remote sensing of environmental phenomena. With motive power from surface-mounted solar arrays or microwave receivers, flight endurance of weeks to months could be anticipated. The proposed system offers unique capability for monitoring oceanic and atmospheric characteristics on local or regional scales. Coastal marine and tropospheric research activities, which require temporal resolutions of 2-72 hours, would be prime application areas. Potential missions might include the monitoring of ocean disposals, episodic marine biological events, and river/ocean interactions. Preliminary sizing and performance calculations are presented along with possible mission scenarios and payload complements.

Youngblood, J. W.↗

NASA NTRS1987

Results of acoustic tests of a Prop-Fan model

Results of acoustic tests in a low speed open jet anechoic wind tunnel are presented for a counter rotation Prop-Fan model. The model tested had 5 front and 5 rear rotor blades with swept planform. Noise spectra are presented showing the influence of operating and configuration variables such as: (1) power absorption, (2) tip speed, (3) rotor-rotor spacing, (4) power split between the front and rear blade rows, (5) variation of the RPM ratio between front and rear blade rows, (6) tractor versus pusher (pylon effects), and (7) angle of attack. In addition to model scale results, calculated levels derived from test are presented showing the influence of the above variables on Effective Perceived Noise Level of a 13.1 ft diameter Prop-Fan at a flyover distance of 1500 ft. It was found that the strongest effects are caused by tip speed and power absorption. A significant finding was that there is an optimum operating tip speed for minimum noise for a given power absorption. Effects of other parametric variations are generally small but measurable. In order to minimize noise to meet airplane certification limits, operation at moderate tip speeds and power absorption is shown to be desirable. Accuracy of predicted Effective Perceived Noise Level is shown to be good with the best accuracy in the 590 to 670 ft/sec tip speed range.

Metzger, F.B.↗

NASA NTRS1988

Measuring Flow With Laser-Speckle Velocimetry

Spatial resolution sufficient for calculation of vorticity.In laser-speckle velocimetry, pulsed or chopped laser beam expanded in one dimension by cylindrical lens to illuminate thin, fan-shaped region of flow measured. Flow seeded by small particles. Lens with optical axis perpendicular to illuminating beam forms image of illuminated particles on photographic plate. Speckle pattern of laser-illuminiated, seeded flow recorded in multiple-exposure photographs and processed to extract data on velocity field. Technique suited for study of vortical flows like those about helicopter rotor blades or airplane wings at high angles of attack.

Smith, C. A.↗

NASA NTRS1944

Application of spring tabs to elevator controls

Equations are presented for calculating the stick-force characteristics obtained with a spring-tab type of elevator control. The main problems encountered in the design of a satisfactory elevator spring tab are to provide stick forces in the desired range, to maintain the force per g sufficiently constant throughout the speed range, to avoid undesirable "feel" of the control in ground handling or in flight at low airspeeds, and to prevent flutter. Examples are presented to show the design features of spring tabs required to solve these problems for airplanes of various sizes.

Phillips, William H↗