Wind field analysis for cantilever loads
Numerical analysis to determine interference effects in local wind flow about Saturn 5 launcher umbilical tower
Ujihara, B. H.↗
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An investigation of the multipath problems associated with the Apollo CSM and LM S-band high gain communication antenna tracking systems
Multipath interference effects on Apollo S band and lunar module steerable antenna tracking system
Osborn, J. D.↗
RTG/science instrument radiation interactions for deep space probes Final technical report, 4 Feb. - 30 Jun. 1969
Interference effects of radiation from plutonium dioxide fueled radioisotope thermoelectric generator
Kaminskas, R. A.↗
Perceptual masking - Peripheral vs central factors
Visual perceptual masking under binocular and dichoptic conditions separating peripheral and central interference effects
Boyle, R. C.↗
Drag of a Supercritical Body of Revolution in Free Flight at Transonic Speeds and Comparison with Wind Tunnel Data
The forebody drag of a supercritical body of revolution was measured in free flight over a Mach number range of 0.85 to 1.05 and a Reynolds number range of 11.5 x 10 to the 6th power to 19.4 x 10 to the 6th power and was compared with wind-tunnel data. The forebody drag coefficient for a Mach number less than 0.96 was 0.111 compared with the wind-tunnel value of 0.103. A gradual increase in the drag occurred in the Langley 8-foot transonic pressure tunnel at a lower Mach number than in the Langley 16-foot transonic tunnel or in the free-flight test. The sharp drag rise occurred near Mach 0.98 in free flight whereas the rise occurred near Mach 0.99 in the Langley 16-foot transonic tunnel. The sharp rise was not as pronounced in the Langley 8-foot transonic pressure tunnel and was probably affected by tunnel-wall-interference effects. The increase occurred more slowly and at a higher Mach number. These results indicate that the drag measurements made in the wind tunnels near Mach 1 were significantly affected by the relative size of the model and the wind tunnel.
Usry, J. W.↗
Aerodynamic characteristics of a six-jet V/STOL configuration with four swing-out lift jets in the transition speed range
A wind-tunnel investigation has been made of the longitudinal aerodynamic characteristics and jet-interference effects of a model of a jet V/STOL variable-sweep fighter airplane that employs four direct-lift engines which swing out from the fuselage and two lift-cruise engines located in the rear part of the fuselage. Data were obtained with two wing areas for various forward speeds and power conditions in the transition speed range. The data are presented without analysis or discussion.
Carter, A. W.↗
Viking configuration pitch damping derivatives as influenced by support interference and test technique at transonic and supersonic speeds.
Experimental studies to determine the 140-deg spherically-blunted cone configuration single degree-of-freedom damping derivative included model sting interference effects at M = 1.76 to 3.0. Minimal sting influences are noted with sting/model diameter ratios of 0.183 and effective sting lengths of 3.55 of the reference length. Increasing the ratio of sting diameter to reference length to 0.53 influences damping considerably at angles of attack from 0 to 3 deg at all Reynolds numbers but shows little or no effect at higher angles. Free-flight model wake geometries (wake neck position and size) correlate well with those from captive model tests with the small diameter sting; this adds confidence to obtaining damping derivatives through captive model testing. The damping characteristics of an early Viking configuration obtained on 11.23 and 4.35% scale models using forced and free oscillation test methods, respectively, in two different test facilities are compared. The agreement is quite good under test conditions.
Steinberg, S.↗
Improved method for aerodynamic analysis of wing-body-tail configurations in subsonic and supersonic flow
Method permits analysis of noncircular bodies and calculation of wing-body interference effects in presence of body closure, two features not previously available. In addition, use of vortex distribution, having linear variation in streamwise direction, results in improved chordwise pressure distributions on wing and tail surfaces.
Fox, C. H. J.↗
V/STOL tilt rotor aircraft study mathematical model for a real time simulation of a tilt rotor aircraft (Boeing Vertol Model 222), volume 8
This report documents the development of a real time mathematical model of a tilt rotor aircraft. This mathematical model is to be used in conjunction with the NASA Flight Simulator for Advanced Aircraft (FSAA) at Ames Research Center for evaluation of aircraft performance and handling qualities. In addition to developing the mathematical model, a parallel programming effort was conducted utilizing Boeing-Vertol's Hybrid Simulation Laboratory for the purpose of developing and evaluating model simplification. The mathematical model is an eleven degree of freedom total force model. This model includes the basic six degree of freedom rigid body .outer loop equations written about the instantaneous center of gravity with the inertial and aerodynamic terms included. The rotor is treated as a point source of forces and moments with appropriate response time lags and actuator dynamics. The wing has one vertical bending and one wing torsion degree of freedom. These structural degrees of freedom are treated on a "quasistatic" basis; i.e., the natural frequencies of vibration of the structure are much higher than the· frequencies of the rigid body motion, and the coupling is in the aerodynamic terms. Each nacelle has an independent pitch degree of freedom about the wing pivot. The aerodynamics of the wing, tail, rotors, landing gear and fuselage are included. Wing and tail mutual interference effects and turbine engine performance and dynamic responses' are represented.
H Rosenstein↗
Canard-wing lift interference related to maneuvering aircraft at subsonic speeds
An investigation was conducted at Mach numbers of 0.7 and 0.9 to determine the lift interference effect of canard location on wing planforms typical of maneuvering fighter configurations. The canard had an exposed area of 16.0 percent of the wing reference area and was located in the plane of the wing or in a position 18.5 percent of the wing mean geometric chord above the wing plane. In addition, the canard could be located at two longitudinal stations. Two different wing planforms were tested: one with a leading-edge sweep angle of 60 deg and the other with a leading-edge sweep angle of 44 deg. The results indicated that although downwash from the canard reduced the wing lift at angles of attack up to approximately 16 deg, the total lift was substantially greater with the canard on than with the canard off. At angles of attack above 16 deg, the canard delayed the wing stall. Changing canard deflection had essentially no effect on the total lift, since the additional lift generated by the canard deflection was lost on the wing due to an increased downwash at the wing from the canard.
Gloss, B. B.↗
Phase change paint tests on Rockwell orbiter/tank and orbiter alone configurations (OH3A/OH3B)
Wind tunnel tests were conducted on scale models of the space shuttle orbiter and external tank. The tests were designed to determine the basic heating rate and interference effects on the orbiter-tank configuration and to analyze the effectiveness of the thermal protective system on the reentry vehicle. The phase change paint techniques were used to determine areodynamic heating rates. Oil flow and schlieren photographs were used for flow visualization.
Quan, M.↗
Aerodynamic investigations into various low speed L/D improvement devices on the 140A/B space shuttle orbiter configuration in the Rockwell International low speed wind tunnel (OA86)
Tests were conducted to investigate various base drag reduction techniques in an attempt to improve Orbiter lift-to-drag ratios and to calculate sting interference effects on the Orbiter aerodynamic characteristics. Test conditions and facilites, and model dimensional data are presented along with the data reduction guidelines and data set/run number collation used for the studies. Aerodynamic force and moment data and the results of stability and control tests are also given.
Mennell, R. C.↗
Theoretical prediction of thick wing and pylon-fuselage-fanpod-nacelle aerodynamic characteristics at subcritical speeds. Part 1: Theory and results
The theoretical development and the comparison of results with data of a thick wing and pylon-fuselage-fanpod-nacelle analysis are presented. The analysis utilizes potential flow theory to compute the surface velocities and pressures, section lift and center of pressure, and the total configuration lift, moment, and vortex drag. The skin friction drag is also estimated in the analysis. The perturbation velocities induced by the wing and pylon, fuselage and fanpod, and nacelle are represented by source and vortex lattices, quadrilateral vortices, and source frustums, respectively. The strengths of these singularities are solved for simultaneously including all interference effects. The wing and pylon planforms, twists, cambers, and thickness distributions, and the fuselage and fanpod geometries can be arbitrary in shape, provided the surface gradients are smooth. The flow through nacelle is assumed to be axisymmetric. An axisymmetric center engine hub can also be included. The pylon and nacelle can be attached to the wing, fuselage, or fanpod.
Tulinius, J. R.↗
Upper-surface blowing nacelle design study for a swept wing airplane at cruise conditions
A study was made to design two types of overwing nacelles for an existing wing-body at a design condition of Mach = 0.8 and C sub L = 0.2. Internal and external surface contours were developed for nacelles having either a D-shaped nozzle or a high-aspect-ratio nozzle for upper-surface blowing in the powered-lift mode of operation. The goal of the design was the development of external nacelle lines that would minimize high-speed aerodynamic interference effects. Each nacelle type was designed for both two- and four-engine airplanes using an iterative process of aerodynamic potential flow analysis. Incremental nacelle drag estimates were made for flow-through wind tunnel models of each configuration.
Gillette, W. B.↗
Study of multiple hologram recording in lithium niobate
The results of detailed experimental and theoretical considerations relating to multiple hologram recording in lithium niobate are reported. The following problem areas are identified and discussed: (1) the angular selectivity of the stored holograms, (2) interference effects due to the crystal surfaces, (3) beam divergence effects, (4) material recording sensitivity, and (5) scattered light from material inhomogeneities.
Gaylord, T. K.↗
Evaluation of tactual displays for flight control
Manual tracking experiments were conducted to determine the suitability of tactual displays for presenting flight-control information in multitask situations. Although tracking error scores are considerably greater than scores obtained with a continuous visual display, preliminary results indicate that inter-task interference effects are substantially less with the tactual display in situations that impose high visual scanning workloads. The single-task performance degradation found with the tactual display appears to be a result of the coding scheme rather than the use of the tactual sensory mode per se. Analysis with the state-variable pilot/vehicle model shows that reliable predictions of tracking errors can be obtained for wide-band tracking systems once the pilot-related model parameters have been adjusted to reflect the pilot-display interaction.
Levison, W. H.↗
Light aircraft lift, drag, and moment prediction: A review and analysis
The historical development of analytical methods for predicting the lift, drag, and pitching moment of complete light aircraft configurations in cruising flight is reviewed. Theoretical methods, based in part on techniques described in the literature and in part on original work, are developed. These methods form the basis for understanding the computer programs given to: (1) compute the lift, drag, and moment of conventional airfoils, (2) extend these two-dimensional characteristics to three dimensions for moderate-to-high aspect ratio unswept wings, (3) plot complete configurations, (4) convert the fuselage geometric data to the correct input format, (5) compute the fuselage lift and drag, (6) compute the lift and moment of symmetrical airfoils to M = 1.0 by a simplified semi-empirical procedure, and (7) compute, in closed form, the pressure distribution over a prolate spheroid at alpha = 0. Comparisons of the predictions with experiment indicate excellent lift and drag agreement for conventional airfoils and wings. Limited comparisons of body-alone drag characteristics yield reasonable agreement. Also included are discussions for interference effects and techniques for summing the results above to obtain predictions for complete configurations.
Smetana, F. O.↗
Heat transfer phase change paint tests of 0.0175-scale models (nos. 21-0 and 46-0) of the Rockwell International space shuttle orbiter in the AEDC tunnel B hypersonic wind tunnel (test OH25A)
Tests were conducted in a hypersonic wind tunnel using various truncated space shuttle orbiter configurations in an attempt to establish the optimum model size for other tests examining body shock-wing leading edge interference effects. The tests were conducted at Mach number 8 using the phase change paint technique. A test description, tabulated data, and tracings of isotherms made from photographs taken during the test are presented.
Dye, W. H.↗