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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.

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At least 199 records · Page 11
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Development of separation due to interaction between a shock wave and a turbulent boundary layer perturbed by rarefaction waves

The interaction of a turbulent boundary layer with rarefaction and shock waves in flows past inclined steps was investigated experimentally and theoretically. The experiments were carried out in supersonic wind tunnels with test sections of 0.6 x 0.6 m and 0.2 x 0.2 m, with adiabatic conditions on the surface model. Based on detailed measurements of pressure fields, velocity, and surface friction, combined with results of optical visualization and analysis of limiting flow lines, three characteristic flow regimes are identified. These are (1) nonseparated flow at small step angles, (2) formation of a local separation zone with a free separation point at moderate step angles, and (3) formation of a large-scale separated flow with a fixed separation point at sufficiently large step angles. The quantitative gasdynamic schemes and surface pressure distributions for these flow regimes are shown for Mach 2.85.

Zheltovodov, A. A.↗
NASA NTRS1994

Comparison of Mixing Calculations for Reacting and Non-Reacting Flows in a Cylindrical Duct

A production 3-D elliptic flow code has been used to calculate non-reacting and reacting flow fields in an experimental mixing section relevant to a rich burn/quick mix/lean burn (RQL) combustion system. A number of test cases have been run to assess the effects of the variation in the number of orifices, mass flow ratio, and rich-zone equivalence ratio on the flow field and mixing rates. The calculated normalized temperature profiles for the non-reacting flow field agree qualitatively well with the normalized conserved variable isopleths for the reacting flow field indicating that non-reacting mixing experiments are appropriate for screening and ranking potential rapid mixing concepts. For a given set of jet momentum-flux ratio, mass flow ratio, and density ratio (J, MR, and DR), the reacting flow calculations show a reduced level of mixing compared to the non-reacting cases. In addition, the rich-zone equivalence ratio has noticeable effect on the mixing flow characteristics for reacting flows.

Oechsle, V. L.↗
NASA NTRS1999

Combustion of Two-Component Miscible Droplets in Reduced Gravity

This paper presents recent theoretical and experimental results from an ongoing research program that deals with reduced-gravity combustion of bi-component droplets initially in the mm size range or larger. The primary objectives of the research are to study the effects of droplet internal flows, thermal and solutal Marangoni stresses, and species volatility differences on liquid species transport and overall combustion phenomena (e.g., gas-phase unsteadiness, burning rates, sooting, radiation, and extinction). The research program utilizes a reduced-gravity environment so that buoyancy effects are rendered negligible. Use of large droplets also facilitates visualization of droplet internal flows, which is important to this research. This program is a continuation of extensive ground based experimental and theoretical research on bi-component droplet combustion which has been ongoing for several years. The focal point of this research program is a flight experiment (Bi-Component Droplet Combustion Experiment, BCDCE). This flight experiment is still under development. However, supporting ground studies have been performed, and preliminary data have been obtained from flight experiments (Fiber Supported Droplet Combustion Experiment, FSDC-1 and FSDC-2). These flight experiments were performed during the STS-73/USML-2 and STS-94/MSL- I missions. In the experiments, droplets composed of low- and high-volatility species are burned. The low-volatility components are initially present in small amounts. As combustion of a droplet proceeds, the liquid surface mass fraction of the low-volatility component will increase with time, resulting in a sudden and temporary decrease in droplet burning rates as the droplet rapidly heats to temperatures close to the boiling point of the low-volatility component. This decrease in burning rates causes a sudden and temporary contraction of the flame. The decrease in burning rates and the flame contraction can be observed experimentally. Measurements of burning rates as well as the onset time for flame contraction allow effective liquid-phase species diffusivities to be calculated, e.g., using asymptotic theory. A goal of the research is to relate effective liquid species diffusivities to droplet internal flow characteristics. Droplet internal flows will be visualized in future flight and ground-based experiments.

Shaw, B. D.↗
NASA NTRS2015

11 Foot Unitary Plan Tunnel Facility Optical Improvement Large Window Analysis

The test section of the 11 by 11-foot Unitary Plan Transonic Wind Tunnel (11-foot UPWT) may receive an upgrade of larger optical windows on both the North and South sides. These new larger windows will provide better access for optical imaging of test article flow phenomena including surface and off body flow characteristics. The installation of these new larger windows will likely produce a change to the aerodynamic characteristics of the flow in the Test Section. In an effort understand the effect of this change, a computational model was employed to predict the flows through the slotted walls, in the test section and around the model before and after the tunnel modification. This report documents the solid CAD model that was created and the inviscid computational analysis that was completed as a preliminary estimate of the effect of the changes.

computer aided design↗
NASA NTRS1995

Aerodynamic Performance and Flow-Field Characteristics of Two Waverider-Derived Hypersonic Cruise Configurations

The component integration of a class of hypersonic high-lift configurations known as waveriders into hypersonic cruise vehicles was evaluated. A wind-tunnel model was developed which integrates realistic vehicle components with two waverider shapes, referred to as the straight-wing and cranked-wing shapes. Both shapes were conical-flow-derived waveriders for a design Mach number of 4.0. Experimental data and limited computational fluid dynamics (CFD) predictions were obtained over a Mach number range of 1.6 to 4.63 at a Reynolds number of 2.0 x 10(exp 6) per foot. The CFD predictions and flow visualization data confirmed the shock attachment characteristics of the baseline waverider shapes and illustrated the waverider flow-field properties. Experimental data showed that no significant performance degradations, in terms of maximum lift-to-drag ratios, occur at off-design Mach numbers for the waverider shapes and the integrated configurations. A comparison of the fully-integrated waverider vehicles to the baseline shapes showed that the performance was significantly degraded when all of the components were added to the waveriders, with the most significant degradation resulting from aftbody closure and the addition of control surfaces. Both fully-integrated configurations were longitudinally unstable over the Mach number range studied with the selected center of gravity location and for unpowered conditions. The cranked-wing configuration provided better lateral-directional stability characteristics than the straight-wing configuration.

Cockrell, Charles E., Jr.↗
NASA NTRS1989

In-flight flow visualization characteristics of the NASA F-18 high alpha research vehicle at high angles of attack

Surface and off-surface flow visualization techniques have been used to visualize the three-dimensional separated flows on the NASA F-18 high alpha research vehicle at high angles of attack. Results near alpha = 25 deg to 26 deg and alpha = 45 deg to 49 deg are presented. Both the forebody and leading-edge extension (LEX) vortex cores and breakdown locations were visualized using smoke. Forebody and LEX vortex separation lines on the surface were defined using an emitted-fluid technique. A laminar separation bubble was also detected on the nose cone using the emitted fluid technique and was similar to that observed in the wind-tunnel test, but not as extensive. Regions of attached, separated, and vortical flow were noted on the wing and the leading-edge flap using tufts and flow cones, and compared well with limited wind-tunnel results.

Fisher, David F.↗
NASA NTRS1991

In-flight flow visualization characteristics of the NASA F-18 high alpha research vehicle at high angles of attack

Surface and off-surface flow visualization techniques were used to visualize the 3-D separated flows on the NASA F-18 high alpha research vehicle at high angles of attack. Results near the alpha = 25 to 26 deg and alpha = 45 to 49 deg are presented. Both the forebody and leading edge extension (LEX) vortex cores and breakdown locations were visualized using smoke. Forebody and LEX vortex separation lines on the surface were defined using an emitted fluid technique. A laminar separation bubble was also detected on the nose cone using the emitted fluid technique and was similar to that observed in the wind tunnel test, but not as extensive. Regions of attached, separated, and vortical flow were noted on the wing and the leading edge flap using tufts and flow cones, and compared well with limited wind tunnel results.

Fisher, David F.↗
NASA NTRS1985

Juncture flow control using leading-edge fillets

Flow measurements have been made in the wake region of a simulated wing-fuselage juncture, with and without leading-edge fillets, in order to assess the effect of leading-edge modifications on the flow field around the juncture. Preliminary results indicate that there is some reduction in juncture drag at moderate angles of attack with the use of leading-edge fillets. There is also evidence of improvement in the flow characteristics on the surface downstream of the juncture when fillets are used. As the fillet size is increased, flow characteristics start deteriorating at some point, and an optimum fillet size may be required to achieve an overall improvement in the flow field.

Kubendran, L. R.↗
NASA NTRS1985

Effect of Steady-State Temperature Distortion on Inlet Flow to a High-Bypass-Ratio Turbofan Engine

The effects of circumferential inlet temperature distortion on the flow characteristics between a distortion generator and a high bypass ratio turbofan engine and through its compression system were evaluated to support the effort to generate analytical models. The flow characteristics are defined by the inlet duct, the flow angles, and the total temperature, total pressure, and static pressure profiles in the inlet duct and through the fan and compressor. The effects of Reynolds number, rotor speed, and distortion extent are also considered.

Soeder, R. H.↗
NASA NTRS1979

Combined pressure and temperature distortion effects on internal flow of a turbofan engine

The flow characteristics obtained experimentally for the compression of a 2-spool TF30-P-3 turbofan engine operating with 180 grad combined pressure and temperature distortion in the inlet flow are presented. The analytical model (Mazzawy and Banks, 1976), 'tuned' during Lewis testing, was used for pretest predictions of the effects that these distortions would have on the engine flow characteristics and the limiting distortion values. The effect of inlet flow distortion on the performance of the engine is discussed, including: (1) the flow between a screen mounted in the inlet duct and the inlet guide vanes; (2) the flow through the compression system; and (3) the effects of the combined distortion and its orientation on the compressor stability limits. It is concluded that the model used in this program was capable of predicting the effects of total pressure, total temperature and combined total pressure-total temperature distortions in terms of flow profiles, inlet flow angles and attenuation of the distortions through the compressor system. It was also capable of predicting the trends of the limiting values experienced with various orientations of the combined distortions.

Braithwaite, W. M.↗
NASA NTRS1972

Results of a study of Mach number and Reynolds number effects on the lee side vortex flow field characteristics of an ogive-cylinder-frustum-cylinder at angles of attack to 25 degrees, volume 1

An experimental program was conducted to survey the lee side vortex flow field about an ogive-cylinder-frustum-cylinder at angles of attack to 25 degrees for two Reynolds numbers at Mach number 0.8, and one Reynolds number at Mach number 1.96. The data were obtained using miniature 5-port conical pressure probes calibrated for angle of attack and roll angle over a Mach number range of 0.6 to 3.0. The results are presented here as local flow field properties and circulation strengths for various body stations.

Foley, J. E.↗
NASA NTRS1975

Improving estimates of streamflow characteristics using LANDSAT-1 (ERTS-1) imagery

The author has identified the following significant results. Streamflow characteristics in the Delmarva Peninsula derived from the records of daily discharge of 20 gaged basins are representative of the full range in flow conditions and include all of those commonly used for design or planning purposes. They include annual flood peaks with recurrence intervals of 2, 5, 10, 25, and 50 years, mean annual discharge, standard deviation of the mean annual discharge, mean monthly discharges, standard deviation of the mean monthly discharges, low-flow characteristics, flood volume characteristics, and the discharge equalled or exceeded 50 percent of the time. Streamflow and basin characteristics were related by a technique of multiple regression using a digital computer. A control group of equations was computed using basin characteristics derived from maps and climatological records. An experimental group of equations was computed using basin characteristics derived from LANDSAT imagery as well as from maps and climatological records. Based on a reduction in standard error of estimate equal to or greater than 10 percent, the equations for 12 stream flow characteristics were substantially improved by adding to the analyses basin characteristics derived from LANDSAT imagery.

Hollyday, E. F.↗
NASA NTRS1999

Flow Quality Measurements in an Aerodynamic Model of NASA Lewis' Icing Research Tunnel

As part of an ongoing effort to improve the aerodynamic flow characteristics of the Icing Research Tunnel (IRT), a modular scale model of the facility was fabricated. This 1/10th-scale model was used to gain further understanding of the flow characteristics in the IRT. The model was outfitted with instrumentation and data acquisition systems to determine pressures, velocities, and flow angles in the settling chamber and test section. Parametric flow quality studies involving the insertion and removal of a model of the IRT's distinctive heat exchanger (cooler) and/or of a honeycomb in the settling chamber were performed. These experiments illustrate the resulting improvement or degradation in flow quality.

Canacci, Victor A.↗
NASA NTRS1983

Inlet, engine, airframe controls integration development for supercruising aircraft

In connection with a consideration of advanced military aircraft systems, attention is given to research for improving the technology of the design of supersonic cruise aircraft. Syberg et al. (1981) have shown that an analytic design method is now available to accurately predict the flow characteristics of axisymmetric supersonic inlets, including off-design angle of attack operation. On the basis of information regarding the inlet flow characteristics, the control system designer can begin the inlet design and development, before wind tunnel testing has begun. The present investigation is concerned with details and status of inlet control technology. A detailed representation of a supersonic propulsion system is developed. This development demonstrates the feasibility of the selected hybrid computational concept.

Houchard, J. H.↗
NASA NTRS1992

Flow characterization in the NASA Ames 16-inch Shock Tunnel

Flow characteristics of NASA's 16-Inch Shock Tunnel are determined for purposes of providing hypersonic propulsion simulation capability. The key tunnel operating parameters are the incident shock speed and reservoir pressure and enthalpy. Flow characteristics of concern are the nozzle exit pressure, temperature, Mach number, Reynolds number, chemical composition, and flow uniformity. Surface mounted gages (for pressure and heat transfer) and nonintrusive optical flow diagnostics (emission and absorption spectroscopy and holographic interferometry) are used to verify tunnel conditions. Experimental measurements are used to validate computational analysis for predicting facility performance, and CFD is used to interpret the free stream optical diagnostic measurements.

Cavolowsky, John A.↗
NASA NTRS1993

In-flight pressure distributions and skin-friction measurements on a subsonic transport high-lift wing section

Flight experiments are being conducted as part of a multiphased subsonic transport high-lift research program for correlation with wind-tunnel and computational results. The NASA Langley Transport Systems Research Vehicle (B737-100 aircraft) is used to obtain in-flight flow characteristics at full-scale Reynolds numbers to contribute to the understanding of 3-D high-lift, multi-element flows including attachment-line transition and relaminarization, confluent boundary-layer development, and flow separation characteristics. Flight test results of pressure distributions and skin friction measurements were obtained for a full-chord wing section including the slat, main-wing, and triple-slotted, Fowler flap elements. Test conditions included a range of flap deflections, chord Reynolds numbers (10 to 21 million), and Mach numbers (0.16 to 0.40). Pressure distributions were obtained at 144 chordwise locations of a wing section (53-percent wing span) using thin pressure belts over the slat, main-wing, and flap elements. Flow characteristics observed in the chordwise pressure distributions included leading-edge regions of high subsonic flows, leading-edge attachment-line locations, slat and main-wing cove-flow separation and reattachment, and trailing-edge flap separation. In addition to the pressure distributions, limited skin-friction measurements were made using Preston-tube probes. Preston-tube measurements on the slat upper surface suggested relaminarization of the turbulent flow introduced by the pressure belt on the slat leading-edge surface when the slat attachment line was laminar. Computational analysis of the in-flight pressure measurements using two-dimensional, viscous multielement methods modified with simple-sweep theory showed reasonable agreement. However, overprediction of the pressures on the flap elements suggests a need for better detailed measurements and improved modeling of confluent boundary layers as well as inclusion of three-dimensional viscous effects in the analysis.

Yip, Long P.↗
NASA NTRS1959

Summary and Analysis of Horizontal-Tail Contribution to Longitudinal Stability of Swept-Wing Airplanes at Low Speeds

Air-flow characteristics behind wings and wing-body combinations are described and are related to the downwash at specific tall locations for unseparated and separated flow conditions. The effects of various parameters and control devices on the air-flow characteristics and tail contribution are analyzed and demonstrated. An attempt has been made to summarize certain data by empirical correlation or theoretical means in a form useful for design. The experimental data herein were obtained mostly at Reynolds numbers greater than 4 x 10(exp 6) and at Mach numbers less than 0.25.

Neely, Robert H.↗