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.
Vortex flows caused by a swept wing are described in some detail. The thin-shear-layer Navier-Stokes equations with an algebraic eddy-viscosity turbulence model are solved. An implicit, factored numerical scheme is used. On the upper surface of the wing, flow is primarily along the free-stream and spanwise direction, respectively, at small and large angles of attack.
The equations are presented for the development of the compressible laminar boundary layer over a yawed infinite cylinder. For compressible flow with a pressure gradient the chordwise and spanwise flows are not independent. Using the Stewartson transformation and a linear viscosity-temperature relation yields a set of three simultaneous ordinary differential equations in a form yielding similar solutions. These equations are solved for stagnation-line flow for surface temperatures from zero to twice the free-stream stagnation temperature and for a wide range of yaw angle and free-stream Mach number. The results indicate that the effect of yaw on the heat-transfer coefficient at the stagnation line depends markedly on the free-stream Mach number. An unusual result of the solutions is that for large yaw angles and stream Mach numbers the chordwise velocity within the boundary layer exceeds the local external chordwise velocity, even for a highly cooled wall.
During a floating zone experiment involving the growth of indium on a recent orbiter mission, (STS 32) oscillation of the zone shapes were observed to occur in response to the background acceleration. An understanding of the nature of the response of the zone shape to forced (g-jitter) oscillations and predictions of its impact on future experiments is of great interest not only to the PI's but to other commercial and academic investigators who plan to fly similar experiments in the orbiter and on space station. Motivated by this, a 15 month study was undertaken to analyze the nature of the g-sensitivity of the STS 32 floating zone crystal growth experiment. Numerical models were used to describe the time-dependent free surface motion of the zone as it responds to the spacecraft residual acceleration. Relevant experimental data concerning the acceleration environment was obtained from the Honeywell in Space Accelerometer (HISA) investigators through MSFC's ACAP program and processed and analyzed. For the indium floating zone experiment, a series of calculations were made using time-dependent axial accelerations g(t). The form of g(t) included simple sinusoidal disturbances as well as actual data (subject to appropriate filtering) measured on the STS 32 mission. Focus was on the calculation of the response of the free surface of the zone as well as the internal flows and internal heat transfer. The influence of solidification on the response of the zone shape was also examined but found to be negligible.
A detailed analysis of the liquid film characteristics and the accompanying heat transfer of a free surface controlled liquid impinging jet onto a rotating disk are presented. The computations were run on a two-dimensional axi-symmetric Eulerian mesh while the free surface was calculated with the volume of fluid method. Flow rates between 3 and 15 1pm with rotational speeds between 50 and 200 rpm are analyzed. The effects of inlet temperature on the film thickness and heat transfer are characterized as well as evaporative effects. The conjugate heating effect is modeled, and was found to effect the heat transfer results the most at both the inner and outer edges of the heated surface. The heat transfer was enhanced with both increasing flow rate and increasing rotational speeds. When evaporative effects were modeled, the evaporation was found to increase the heat transfer at the lower flow rates the most because of a fully developed thermal field that was achieved. The evaporative effects did not significantly enhance the heat transfer at the higher flow rates.
Utilizing geometry based static phase separation exhibited in the radial vaned capillary drying tray, a system was conceived to recover water from brine. This technology has been named the Capillary BRIC; abbreviated CapiBRIC. The CapiBRIC utilizes a capillary drying tray within a drying chamber. Water is recovered from clean water vapor evaporating from the free surface leaving waste brine solids behind. A novel approach of optimizing the containment geometry to support passive capillary flow and static phase separation provides the opportunity for a low power system that is not as susceptible to fouling as membranes or other technologies employing physical barriers across the free brine surface to achieve phase separation in microgravity. Having been optimized for operation in microgravity, full-scale testing of the CapiBRIC as designed cannot be performed on the ground as the force of gravity would dominate over the capillary forces. However, subscale units relevant to full-scale design were used to characterize fill rates, containment stability, and interaction with a variable volume reservoir in the PSU Dryden Drop Tower (DDT) facility. PSU also using tested units scaled such that capillary forces dominated in a 1-g environment to characterize evaporation from a free-surface in 1-g upward, sideways and downward orientations. In order to augment the subscale testing performed by PSU, a full scale 1-g analogue of the CapiBRIC drying unit was initiated to help validate performance predictions regarding expected water recovery ratio, estimated processing time, and interface definitions for inlets, outlets, and internal processes, including vent gas composition. This paper describes the design, development and test of the terrestrial CapiBRIC prototypes.
A simple analytical method has been introduced for aerohydrodynamic load analysis of an airborne configuration during water ditching. The method employs an aerodynamic panel code, based on linear potential flow theory, to simulate the flow of air and water around an aircraft configuration. The free surface separating the air and water region is represented by doublet sheet singularities. Although all the theoretical load distributions are computed for air, provisions are made to correct the pressure coefficients obtained on the configuration wetted surfaces to account for the water density. As an analytical tool, the Vortex Separation Aerodynamic (VSAERO) code is chosen to carry out the present investigation. After assessing the validity of the method, its first application is to analyze the water ditching of the Space Shuttle configuration at a 12 degree attitude.
A review of natural laminar flow (NLF) flight experiences over the period from the 1930's to the present has been given to provide information on the achievability and maintainability of NLF in typical airplane operating environments. Significant effects of loss of laminar flow on airplane performance have been observed for several airplanes, indicating the importance of providing information on these changes to laminar flow airplane operators. Significant changes in airplane stability and control and maximum lift were observed in flight experiments with the loss of laminar flow. However, these effects can be avoided by proper selection of airfoils. Conservative laminar flow airfoil designs should be employed which do not experience significant loss of lift (caused by flow separation) upon the loss of laminar flow. Mechanisms have been observed for the effects of insect accumulation, flight through clouds and precipitation, and propeller slipstreams on laminar flow behavior. Fixed transition testing, in addition to free transition testing, is recommended as a new standard procedure for airplanes with surfaces designed to support laminar flow.
A simple analytical method has been introduced for aerohydrodynamic load analysis of an airborne configuration during water ditching. The method employs an aerodynamic panel code, based on linear potential flow theory, to simulate the flow of air and water around an aircraft configuration. The free surface separating the air and water region is represented by doublet" singularities. Although all the theoretical load distributions are computed for air, provisions are made to correct the pressure coefficients obtained on the configuration-wetted surfaces to account for the water density. As an analytical tool, a vortex aerodynamic code is chosen to carry out the present investigation. After assessing· the validity of the method, an application is presented for the water ditching of the Space Shuttle configuration at a 12-deg attitude.
The properties of 2-D shock wave/turbulent boundary layer interaction flows were calculated by using a compressible turbulent Navier-Stokes numerical computational code. Interaction flows caused by oblique shock wave impingement on the turbulent boundary layer flow were considered. The oblique shock waves were induced with shock generators at angles of attack less than 10 degs in supersonic flows. The surface temperatures were kept at near-adiabatic (ratio of wall static temperature to free stream total temperature) and cold wall (ratio of wall static temperature to free stream total temperature) conditions. The computational results were studied for the surface heat transfer, velocity temperature correlation, and turbulent shear stress in the interaction flow fields. Comparisons of the computational results with existing measurements indicated that (1) the surface heat transfer rates and surface pressures could be correlated with Holden's relationship, (2) the mean flow streamwise velocity components and static temperatures could be correlated with Crocco's relationship if flow separation did not occur, and (3) the Baldwin-Lomax turbulence model should be modified for turbulent shear stress computations in the interaction flows.
A simple analytical method has been introduced for aerohydrodynamic load analysis of an airborne configuration during water ditching. The method employs an aerodynamic panel code, based on linear potential flow theory, to simulate the flow of air and water around an aircraft configuration. The free surface separating the air and water region is represented by doublet sheet singularities. Although all the theoretical load distributions are computed for air, provisions are made to correct the pressure coefficients obtained on the configuration wetted surfaces to account for the water density. As an analytical tool, the Vortex Separation Aerodynamic (VSAERO) code is chosen to carry out the present investigation. After assessing the validity of the method, its first application is to analyze the water ditching of the Space Shuttle configuration at a 12 degree attitude.
Short duration, cold flow test technique which duplicates pressure effects of underexpanded free hot rocket exhaust jets impinging upon adjacent surfaces
A computational method for designing shock-free, quasi-three-dimensional, transonic, turbomachinery blades is described. Shock-free designs are found by implementing Sobieczky's fictitious gas principle in the analysis of a baseline shape, resulting in an elliptic solution that is incorrect in the supersonic domain. Shock-free designs are obtained by combining the subsonic portion of this solution with a characteristic calculation of the correct supersonic flow using the sonic line data from the fictitious elliptic solution. This provides a new, shock-free blade design. Examples presented include the removal of shocks from two blades in quasi-three-dimensional flow and the development of a series of shock-free two-dimensional stators. The new designs all include modifications to the upper surface of an experimental stator blade developed at NASA Lewis Research Center. While the designs presented here are for inviscid flow, the same concepts have been successfully applied to the shock-free design of airfoils and three-dimensional wings with viscous effects. The extension of the present method to viscous flows is straightforward given a suitable analysis algorithm for the flow.
An experimental study of oscillatory thermocapillary in small cylindrical containers with a heating wire placed along the center axis is performed by investigating the flow structures and temperature distributions under various conditions. To supplement the flow visualization the surface is scanned using an infrared imager. Here, 2 cS viscosity (Pr = 27) silicone oil is used as the test fluid. It is observed that beyond a certain temperature difference between the container wall and the heating wire, a distinctive unsteady flow pattern appears. This unsteady phenomenon is identified as oscillatory thermocapillary. After the onset of oscillations the flow structure becomes nonaxisymmetric and wave motion is observed at the free surface. It is shown that the critical temperature difference is independent of container dimensions if the aspect ratio is fixed.
A series of experiments was conducted on flow fields about two airfoil models whose sections are slight modifications of the original Whitcomb supercritical airfoil section. Data obtained include surface static-pressure distributions, far-wake surveys, oil-flow photographs, pitot-pressure surveys in the viscous regions, and holographic interferograms. These data were obtained for different combinations of lift coefficient and free-stream Mach number, which included both subcritical cases and flows with upper-surface shock waves. The availability of both pitot-pressure data and density data from interferograms allowed determination of flow-field properties in the vicinity of the trailing edge and in the wake without recourse to any assumptions about the local static pressure. The data show that significant static-pressure gradients normal to viscous layers exist in this region, and that they persist to approximately 10% chord downstream of the trailing edge. Comparisons are made between measured boundary-layer properties and results from boundary-layer computations that employed measured static-pressure distributions, as well as comparisons between data and results of airfoil flow-field computations.
Wind-tunnel tests were conducted on an ogive-cylinder model with two axisymmetric protuberances having cone frustum angles of cone = 23 deg and 45 deg that were used to generate detached shock waves and the resulting separated flow areas downstream of the shock. The tests were conducted in a 9 by 7 foot supersonic wind tunnel at a free-stream Mach number of 2.0 and at Reynolds numbers of 1.5 x 1 million and 3.9 x 1 million, based on body diameter. The model had an afterbody fineness ratio of 8.3, and the ogive nose had a fineness ratio of 3.0. Two characteristics of the fluctuating pressures in surface vortex flows that result from the crossflow component, (velocity along the tunnel longitudinal axis free stream angle of attack), in combination with changes in the longitudinal pressure gradient were measured: (1) the broadband, rms-pressure coefficients and (2) the power spectral densities. Measurements are presented for various flow regions on the model such as the attached turbulent boundary layer, the detached frustum shock wave, and separated flow areas. The results indicate that the pressure fluctuations around or in the neighborhood of the foci of the vortex flows had broadband intensities and power spectral densities nearly identical to the levels previously measured in separated-flow regions at angles of attack of 0 deg.
The flow over wave shaped surfaces is studied using a Navier Stokes solver. Detailed comparisons with theoretical results are presented, including the stability of a laminar flow over wavy surfaces. Drag characteristics of nonplanar surfaces are predicted using the Navier-Stokes solver. The secondary instabilities of wall bounded and free shear flows are also discussed.
Results of computer code time dependent solutions of the two dimensional compressible Navier-Stokes equations and the results of independent experiments are compared to verify the Mach number range for instabilities in the transonic flow field about a 14 percent thick biconvex airfoil at an angle of attack of 0 deg and a Reynolds number of 7 million. The experiments were conducted in a transonic, slotted wall wind tunnel. The computer code included an algebraic eddy viscosity turbulence model developed for steady flows, and all computations were made using free flight boundary conditions. All of the features documented experimentally for both steady and unsteady flows were predicted qualitatively; even with the above simplifications, the predictions were, on the whole, in good quantitative agreement with experiment. In particular, predicted time histories of shock wave position, surface pressures, lift, and pitching moment were found to be in very good agreement with experiment for an unsteady flow. Depending upon the free stream Mach number for steady flows, the surface pressure downstream of the shock wave or the shock wave location was not well predicted.
A description is presented of the development of the boundary layer on the lip and diffuser surface of a subsonic inlet at arbitrary operating conditions of mass flow rate, free stream velocity and incidence angle. Both laminar separation on the lip and turbulent separation in the diffuser are discussed. The agreement of the theoretical results with model experimental data illustrates the capability of the theory to predict separation. The effects of throat Mach number, inlet size, and surface roughness on boundary layer development and separation are illustrated.