SEARCH · Engineering Papers
Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
The experimental methodologies are illustrated by graphs, charts and line drawings. Typical ultrasonic echo signals for dry and wet ice growth, ice accretion rates for various tunnel configurations, the experimental configuration for flight tests of the ultrasonic measuring system and heat balance models used to predict ice growth are among the topics that are illustrated and briefly discussed.
The ability of the Vortex Separation AEROdynamics (VSAERO) program to calculate aerodynamic loads on wings due to interaction with free vortices was studied. The loads were calculated for various positions of a downstream following wing relative to an upstream vortex-generating wing. Calculated vortex-induced span loads, rolling-moment coefficients, and lift coefficients on the following wing were compared with experimental results of McMillan et al. and El-Ramly et al. Comparisons of calculated and experimental vortex tangential velocities were also made.
The theoretical and experimental status of helicopter impulsive noise is reviewed. The two major source mechanisms of helicopter impulsive noise are addressed: high-speed impulsive noise and blade-vortex interaction impulsive noise. A thorough physical explanation of both generating mechanisms is presented together with model and full-scale measurements of the phenomena. Current theoretical prediction methods are compared with experimental findings of isolated rotor tests. The noise generating mechanisms of high speed impulsive noise are fairly well understood - theory and experiment compare nicely over Mach number ranges typical of today's helicopters. For the case of blade-vortex interaction noise, understanding of noise generating mechanisms and theoretical comparison with experiment are less satisfactory. Several methods for improving theory/experiment are suggested.
A theoretical and experimental investigation of turbulent flows inside a 180-deg strongly curved channel is included in the present study. A 10 by 100 centimeter water channel with a 10 centimeter radius turn was employed for the experimental study. Data measured are velocity profiles and distributions of longitudinal turbulence intensity. Numerical investigations are performed by solving the incompressible Navier-Stokes equations and turbulence models using a finite difference flow solver in generalized curvilinear coordinate systems. A widely used two-equation turbulence model and its modified version, which includes the effect of strong curvature on the turbulence length scale through a relation of algebraic Reynolds stress model, are employed to predict the turbulence quantities inside the flow field. Improvements on the predictions are obtained using the present approach.
Experimental data were obtained for the deicing characteristics of a stationary UH-1H helicopter blade which had been fitted with an electrothermal deicer assembly. The tests were run in the NASA Lewis Icing Research Tunnel, and yielded transient temperature responses for the substrate, heater and abrasion shield at selected positions around the blade. The data at the abrasion shield-ice interface clearly documented when melting, shedding or refreezing occurred. Comparisons were made between the experimental data and a one-dimensional numerical model. The agreement was generally very good, with the simulations being shown to be capable of predicting the transient temperature responses along with phase change and ice shedding. At many blade positions, the model was capable of accurately simulating the thermal response of the electrothermal deicer assembly.
The theoretical and experimental status of helicopter impulsive noise is reviewed. The two major source mechanisms of helicopter impulsive noise are addressed: high-speed impulsive noise and blade-vortex interaction impulsive noise. A thorough physical explanation of both generating mechanism is presented together with model and full-scale measurements of the phenomena. Current theoretical prediction methods are compared with experimental findings of isolated rotor tests. The noise generating mechanism of high speed impulsive noise are fairly well understood - theory and experiment compare nicely over Mach number ranges typical of today's helicopters. For the case of blade-vortex interaction noise, understanding of noise generating mechanisms and theoretical comparison with experiment are less satisfactory. Several methods for improving theory-experiment are suggested.
This paper presents a computational and experimental investigation of supersonic flow past a cavity in a flat plate. The source of the particular interest in this problem is the ongoing study of the aerodynamic interference effects between a separating store and its bay in the parent body. An upwind relaxation scheme, utilizing flux vector splitting and line-Gauss-Seidel iterations, is used to solve Reynolds-averaged Navier-Stokes equations. Spatial discretizations of this two-dimensional analysis are based on implicit and finite-volume methods. Turbulence is modeled and shocks are captured. The flowfield of the symmetry plane at the half-width is computationally visualized and all flow properties are computed. Experimental tests are conducted in the Langley Unitary Plan Wind Tunnel to measure wall pressures and to capture schlieren photographs. Qualitative as well as quantitative data of computations and experiments agree very well. These two vehicles of investigation are merged to show open, closed and transitional cavity flow behaviors.
An experimental method has been developed to determine the droplet impingement characteristics on two- and three-dimensional bodies. The experimental results provide the essential droplet impingement data required to validate particle trajectory codes, used in aircraft icing analyses and engine inlet particle separator analyses. A body whose water droplet impingement characteristics are required is covered at strategic locations by thin strips of moisture absorbing (blotter) paper, and then exposed to an air stream containing a dyed-water spray cloud. Water droplet impingement data are extracted from the dyed blotter strips, by measuring the optical reflectance of the dye deposit on the strips, using an automated reflectometer. Impingement efficiency data obtained for a NACA 65(2)015 airfoil section, a supercritical airfoil section, and Being 737-300 and axisymmetric inlet models are presented in this paper.
The performance of a fixed-geometry, swept, mixed compression hypersonic inlet is presented. The experimental evaluation was conducted for a Mach number of 6.0 and for several angles of attack. The measured surface pressures and pitot pressure surveys at the inlet throat are compared to computations using a three-dimensional Euler code and an integral boundary layer theory. Unique features of the intake design, including the boundary layer control, insure a high inlet performance. The experimental data show the inlet has a high mass averaged total pressure recovery, a high mass capture and nearly uniform flow diffusion. The swept inlet exhibits excellent starting characteristics, and high flow stability at angle of attack.
An experimental study was conducted to determine the fracture behavior of centernotched, unidirectional graphite/epoxy laminates when subjected to tensile loading. The actual behavior is compared to the behavior predicted by a mathematical model based on classical shear-lag assumptions. Experimental results show the fracture behavior to consist of two modes of longitudinal matrix splitting. After split initiation, a region of slow, stable split growth occurs. This is followed by a distinct transition into a rapid split growth region. However, the shear-lag model predicts a single region of rapid longitudinal split growth behavior with unbounded splitting occurring after approximately a ten percent increase in the applied stress over the split initiation stress. It is therefore concluded that the classical shear-lag model does not accurately predict the complete longitudinal split growth behavior in a brittle matrix composite. This shortcoming in the model and some results obtained from related elasticity solutions are discussed and it is suggested that the slow, stable split growth region is likely dominated by transverse matrix normal stresses which are not accounted for in the classical shear-lag model.
An analytical and experimental investigation of the effects of mistuning on propfan subsonic flutter was performed. The analytical model is based on the normal modes of a rotating composite blade and a three-dimensional subsonic unsteady lifting surface aerodynamic theory. Theoretical and experimental results are compared for selected cases at different blade pitch angles, rotational speeds, and free-stream Mach numbers. The comparison shows a reasonably good agreement between theory and experiment. Both theory and experiment showed that combined mode shape, frequency, and aerodynamic mistuning can have a beneficial or adverse effect on blade damping depending on Mach number. Additional parametric results showed that alternative blade frequency mistuning does not have enough potential for it to be used as a passive flutter control in propfans similar to the one studied. It can be inferred from the results that a laminated composite propfan blade can be tailored to optimize its flutter speed by selecting the proper ply angles.
The results of a three-dimensional numerical simulation of Boussinesq free convection in a horizontal differentially heated cylinder are presented. The computation was based on a Samarskii-Andreyev scheme (described by Leong, 1981) and a false-transient advancement in time, with vorticity, velocity, and temperature as dependent variables. Solutions for velocity and temperature distributions were obtained for Rayleigh numbers (based on the radius) Ra = 74-18,700, thus covering the core- and boundary-layer-driven regimes. Numerical solutions are compared with asymptotic analytical solutions and experimental data. The numerical results well represent the complex three-dimensional flows found experimentally.
The Joint Army, Navy, NASA, Air Force (JANNAF) rocket-engine performance-prediction procedure is based on the use of various reference computer programs. One of the reference programs for nozzle analysis is the Two-Dimensional Kinetics (TDK) Program. The purpose of this report is to calibrate the JANNAF procedure that has been incorporated into the December 1984 version of the TDK program for the high-area-ratio rocket-engine regime. The calibration was accomplished by modeling the performance of a 1030:1 rocket nozzle tested at NASA Lewis. A detailed description of the test conditions and TDK input parameters is given. The reuslts indicate that the computer code predicts delivered vacuum specific impulse to within 0.12 to 1.9 percent of the experimental data. Vacuum thrust coefficient predictions were within + or - 1.3 percent of experimental results. Predictions of wall static pressure were within approximately + or - 5 percent of the measured values.
Experimental longitudinal and lateral-directional aerodynamics were obtained for a generic aerodynamics were obtaiend for a generic winged-cone configuration having possible application as a transatmospheric vehicle concept. Data were obtained at Mach numbers from 0.6 to 20.0; Reynolds numbers, based on model length, between 2.5 and 5.3 million; and angles of attack from -4 to 20 deg. Results indicate a longitudinal center-of-pressure travel of about 23 percent of the fuselage length for the test Mach number range, with longitudinal instabilities noted at high-supersonic to hypersonic Mach numbers. These instabilities are coupled with directional instability at similar Mach numbers. Predictions with analytic codes, namely, the USAF DATCOM and the tangent-cone option of the Hypersonic Arbitrary Body Program, provided fair agreement with the experimental aerodynamic characteristics at low angles-of-attack.
In this paper, results from two experimental and two computational investigations of plane turbulent mixing layers are presented and compared. The experimental results include flow visualization data using the smoke laser technique and mean flow and turbulence measurements obtained with hot X-wires and a two-component LDV. Reasonably good agreement is found among these techniques, at least for turbulence quantities up to second order. Reynolds-averaged computations are successful at capturing the complete evolution of the mixing layer, including wake effects in the near field and approach to self-preservation in the far field. The two-dimensional vortex method shows excellent qualitative and quantitative agreement with measured data for the forced mixing layer. For the unforced layer, the results seem to indicate that a three-dimensional computation may be necessary.
The recovery of polished 600-micron-thick 7-mm-diameter disk samples and powder samples of the Leedey L6 chondrite from 29-70-GPa experimental shocks is investigated experimentally under vacuum conditions, with a focus on the possible direct shock loss of radiogenic Ar. The results of petrographic examination of the shocked samples and Ar isotope measurements are presented in tables, graphs, and micrographs and characterized in detail. The principal damage mechanisms are identified as mechanical disaggregation and increasing cataclasis, with little melting. The degassing of radiogenic Ar is found to be relatively modest and to increase with shock pressure and sample porosity; it is attributed to a combination of mechanical grain disruption and feldspar melting. It is inferred that Ar loss in naturally shocked materials (especially those showing few signs of melting) probably occurs during extended exposure to high temperatures during impact formation.
Experimental Assembly of Structures in EVA (EASE) objectives, experimental protocol, neutral buoyancy simulation, task time distribution, assembly task performance, metabolic rate/biomedical readouts are summarized. This presentation is shown in charts, figures, and graphs.
The 2-D and 3-D advance computer codes being developed for use in the design of such hypersonic aircraft as the National Aero-Space Plane require comparison of the computational results with a broad spectrum of experimental data to fully assess the validity of the codes. This is particularly true for complex flow fields with control surfaces present and for flows with separation, such as leeside flow. Therefore, the objective is to provide a hypersonic experimental data base required for validation of advanced computational fluid dynamics (CFD) computer codes and for development of more thorough understanding of the flow physics necessary for these codes. This is being done by implementing a comprehensive test program for a generic all-body hypersonic aircraft model in the NASA/Ames 3.5 foot Hypersonic Wind Tunnel over a broad range of test conditions to obtain pertinent surface and flowfield data. Results from the flow visualization portion of the investigation are presented.