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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 631 records · Page 35
All sourcesNASA NTRSDOE OSTI
NASA NTRS2011

An Examination of the Effect of Boundary Layer Ingestion on Turboelectric Distributed Propulsion Systems

A Turboelectric Distributed Propulsion (TeDP) system differs from other propulsion systems by the use of electrical power to transmit power from the turbine to the fan. Electrical power can be efficiently transmitted over longer distances and with complex topologies. Also the use of power inverters allows the generator and motors speeds to be independent of one another. This decoupling allows the aircraft designer to place the core engines and the fans in locations most advantageous for each. The result can be very different installation environments for the different devices. Thus the installation effects on this system can be quite different than conventional turbofans where the fan and core both see the same installed environments. This paper examines a propulsion system consisting of two superconducting generators, each driven by a turboshaft engine located so that their inlets ingest freestream air, superconducting electrical transmission lines, and an array of superconducting motor driven fan positioned across the upper/rear fuselage area of a hybrid wing body aircraft in a continuous nacelle that ingests all of the upper fuselage boundary layer. The effect of ingesting the boundary layer on the design of the system with a range of design pressure ratios is examined. Also the impact of ingesting the boundary layer on off-design performance is examined. The results show that when examining different design fan pressure ratios it is important to recalculate of the boundary layer mass-average Pt and MN up the height for each inlet height during convergence of the design point for each fan design pressure ratio examined. Correct estimation of off-design performance is dependent on the height of the column of air measured from the aircraft surface immediately prior to any external diffusion that will flow through the fan propulsors. The mass-averaged Pt and MN calculated for this column of air determine the Pt and MN seen by the propulsor inlet. Since the height of this column will change as the amount of air passing through the fans change as the propulsion system is throttled, and since the mass-average Pt and MN varies by height, this capture height must be recalculated as the airflow through the propulsor is varied as the off-design performance point is converged.

Felder, James L.↗
NASA NTRS1977

Effects of jet decay rate on jet induced loads on a flat plate

Experimental modelling of the interaction between a jet and an aircraft wing or fuselage in VTOL aircraft was undertaken using a cold jet exiting perpendicular to a flat plate in a uniform cross-flow. Effects of jet decay rate and jet-to-cross-flow velocity ratio, R, on the induced load distribution were investigated. Jet decay rate was increased by using cylindrical centerbodies submerged in the jet nozzle, which caused nonuniform initial jet velocity profiles. Quicker jet decay rate, corresponding to the presence of a centerbody, resulted in as much as 50% reduction in the induced pressure loads on the plate. This has implications in interpretation of results from earlier VTOL model studies of jet induced loads, where the jets have often had relatively slow decay rates due to uniform initial velocity profiles

Kuhlman, J. M.↗
NASA NTRS1979

Longitudinal aerodynamic characteristics of a vectored-engine-over-wing configuration at subsonic speeds

The Langley V/STOL tunnel was used to determine the effects of vectoring exhaust flow on the longitudinal aerodynamic characteristics of a vectored-engine-over-wing configuration. Vectoring was accomplished by blowing from over-wing-mounted engines over a variable trailing-edge flap. Effects of varying canard geometry and wing leading-edge geometry were investigated. Wind-tunnel data were obtained at a Mach number of 0.186 for an angle-of-attack range from -20 deg to 24 deg and engine nozzle pressure ratios from 1.0 (jet off) to approximately 3.75.

Leavitt, L. D.↗
NASA NTRS1980

Preliminary study of advanced turboprop and turboshaft engines for light aircraft

The effects of engine configuration, advanced component technology, compressor pressure ratio and turbine rotor-inlet temperature on such figures of merit as vehicle gross weight, mission fuel, aircraft acquisition cost, operating, cost and life cycle cost are determined for three fixed- and two rotary-wing aircraft. Compared with a current production turboprop, an advanced technology (1988) engine results in a 23 percent decrease in specific fuel consumption. Depending on the figure of merit and the mission, turbine engine cost reductions required to achieve aircraft cost parity with a current spark ignition reciprocating (SIR) engine vary from 0 to 60 percent and from 6 to 74 percent with a hypothetical advanced SIR engine. Compared with a hypothetical turboshaft using currently available technology (1978), an advanced technology (1988) engine installed in a light twin-engine helicopter results in a 16 percent reduction in mission fuel and about 11 percent in most of the other figures of merit.

Knip, G.↗
NASA NTRS1980

Spectrophotometry of H II regions and the nucleus of NGC 1566

The paper presents spectrophotometric observations of the nucleus and nine H II regions in the spiral arms of the Sbc galaxy NGC 1566. The emission line spectra of the H II regions are characterized by a low grade of ionization, and most of the line intensities do not show radial gradients; the N II/O II forbidden line ratio appears to increase toward the nucleus, and the forbidden O III/H-beta ratio may decrease in the same direction. This is explained by the increase of the oxygen abundance and decrease of the electron temperature toward the nucleus; the spectra confirm the presence of the weak, broad emission wings at H-alpha, and the nuclear spectrum of NGC 1566 resembles that of a giant H II region rather than that of a Seyfert galaxy.

Hawley, S. A.↗
NASA NTRS1981

Density sensitive C II lines in cool stars of low gravity

It is shown that the relative intensities of emission lines within the multiplet UV 0.01 of C II, around 2325 A, are sensitive to electron density in the range 10 to the 9th to 10 to the 7th per cu cm. The lines therefore offer a valuable method for measuring electron densities in the chromospheres of late-type giants and supergiants. Calculated line ratios are compared with those observed in a range of objects.

Stencel, R. E.↗
NASA NTRS1975

Approximate method for calculating transonic flow about lifting wing-body configurations: Computer program and user's manual

A user's manual is presented for a computer program which calculates inviscid flow about lifting configurations in the free-stream Mach-number range from zero to low supersonic. Angles of attack of the order of the configuration thickness-length ratio and less can be calculated. An approximate formulation was used which accounts for shock waves, leading-edge separation and wind-tunnel wall effects.

Barnwell, R. W.↗
DOE OSTI2018

Multi-depth suspended sediment estimation using high-resolution remote-sensing UAV in Maumee River, Ohio

Satellite remote-sensing has been widely used to map suspended sediment concentration (SSC) in waterbodies. Current development of the unmanned aerial vehicle (UAV) technology allows mapping of SSC at finer spatial resolution providing high flexibility in terms of cost and acquisition time. However, the technology is immature and transfer of empirical algorithms from existing remote-sensing technologies to UAV still has to be explored. Here, this study uses the MicaSense Sequoia sensor with four bands (green, red, red edge, and near-infrared [NIR]) mounted on-board a fixed-wing UAV to map SSC within the Maumee River in Ohio, USA, at multiple depth intervals (15, 61, 91, and 182 cm). The simple linear and stepwise regression models show the advantage of multiple bands and band ratios over single bands in mapping SSC. The findings show a limited performance of the Sequoia sensor when compared to field spectroradiometer measurements. In all cases but one, the adjusted coefficient of determination (R$2\atop{adj}$) values is lower for the UAV data. The regression equations become similar at and below a depth of 0–61 cm, and R$2\atop{adj}$ become constant at and below a depth of 0–91 cm. While the spectroradiometer-related equations are sensitive to a wider spectral range (from green at the surface to NIR wavelength at 182 cm depth), the UAV-related equations are insensitive to green spectrum and they include a narrower spectral range (from red to NIR) over all depth increments. Field spectroradiometer measurements exhibit a strong relationship with cumulative SSC at 182 cm depth (0–182 cm) (R$2\atop{adj}$ = 0.71) whereas UAV reflectance data show the best relationship with SSC at 91 cm (0–91 cm) (R$2\atop{adj}$ = 0.56) suggesting that ~91 cm may be an optimal depth for UAV under given conditions. In conclusion, the results show that UAVs can be a practical but somewhat limited tool to monitor SSC in small- to medium-sized rivers.

47 OTHER INSTRUMENTATION↗
NASA NTRS1997

Vortex leading edge flap assembly for supersonic airplanes

A leading edge flap (16) for supersonic transport airplanes is disclosed. In its stowed position, the leading edge flap forms the lower surface of the wing leading edge up to the horizontal center of the leading edge radius. For low speed operation, the vortex leading edge flap moves forward and rotates down. The upward curve of the flap leading edge triggers flow separation on the flap and rotational flow on the upper surface of the flap (vortex). The rounded shape of the upper fixed leading edge provides the conditions for a controlled reattachment of the flow on the upper wing surface and therefore a stable vortex. The vortex generates lift and a nose-up pitching moment. This improves maximum lift at low speed, reduces attitude for a given lift coefficient and improves lift to drag ratio. The mechanism (27) to move the vortex flap consists of two spanwise supports (24) with two diverging straight tracks (64 and 68) each and a screw drive mechanism (62) in the center of the flap panel (29). The flap motion is essentially normal to the airloads and therefore requires only low actuation forces.

Rudolph, Peter K. C.↗
NASA NTRS2010

Fan and Open-Rotor Noise

This presentation is a technical progress report and near term outlook for work on fan (in-duct) and open-rotor (high speed propeller) noise funded by NASA's Fundamental Aeronautics Program, Subsonic Fixed Wing (SFW) Project and the Integrated Systems Research Program, Environmentally Responsible Aircraft Project. Sections of the presentation cover: the system level metrics are outlined for the SFW timeframes (2015, 2020 1 2025); the Ultra-High Bypass ratio technology development roadmap; a feasibility study for a low technology readiness level fan test rig; the development plan for a turbomachinery oriented computational aero-acoustics code; and systems analysis work on open-rotor modeling.

Miller, Christopher J.↗
NASA NTRS2013

NASA Fixed Wing Project Propulsion Research and Technology Development Activities to Reduce Thrust Specific Energy Consumption

This paper presents an overview of the propulsion research and technology portfolio of NASA Fundamental Aeronautics Program Fixed Wing Project. The research is aimed at significantly reducing the thrust specific fuel/energy consumption of notional advanced fixed wing aircraft (by 60 % relative to a baseline Boeing 737-800 aircraft with CFM56-7B engines) in the 2030-2035 time frame. The research investments described herein are aimed at improving propulsive efficiency through higher bypass ratio fans, improving thermal efficiency through compact high overall pressure ratio gas generators, and exploring the potential benefits of boundary layer ingestion propulsion and hybrid gas-electric propulsion concepts.

Hathaway, Michael D.↗
NASA NTRS2014

Spectroscopic Measurement Techniques for Aerospace Flows

The conditions that characterize aerospace flows are so varied, that a single diagnostic technique is not sufficient for its measurement. Fluid dynamists use knowledge of similarity to help categorize and focus on different flow conditions. For example, the Reynolds number represents the ratio of inertial to viscous forces in a flow. When the velocity scales, length scales, and gas density are large and the magnitude of the molecular viscosity is low, the Reynolds number becomes large. This corresponds to large scale vehicles (e.g Airbus A380), fast moving objects (e.g. artillery projectiles), vehicles in dense fluids (e.g. submarine in water), or flows with low dynamic viscosity (e.g. skydiver in air). In each of these cases, the inertial forces dominate viscous forces, and unsteady turbulent fluctuations in the flow variables are observed. In contrast, flows with small length scales (e.g. dispersion of micro-particles in a solid rocket nozzle), slow moving objects (e.g. micro aerial vehicles), flows with low density gases (e.g. atmospheric re-entry), or fluids with a large magnitude of viscosity (e.g. engine coolant flow), all have low Reynolds numbers. In these cases, viscous forces become very important and often the flows can be steady and laminar. The Mach number, which is the ratio of the velocity to the speed of sound in the medium, also helps to differentiate types of flows. At very low Mach numbers, acoustic waves travel much faster than the object, and the flow can be assumed to be incompressible (e.g. Cessna 172 aircraft). As the object speed approaches the speed of sound, the gas density can become variable (e.g. flow over wing of Learjet 85). When the object speed is higher than the speed of sound (Ma > 1), the presences of shock waves and other gas dynamic features can become important to the vehicle performance (e.g. SR-71 Blackbird). In the hypersonic flow regime (Ma > 5), large changes in temperature begin to affect flow properties, causing real-gas effects to occur (e.g. X-43 Scramjet). At even higher Mach numbers, chemistry and nonequilibrium effects come into play (e.g. Startdust re-entry capsule), further complicating the measurement. These limits can be predicted by calculating the ratio of chemical and thermal relaxation time to the flow time scales. Other non-dimensional numbers can be used to further differentiate types of aerospace flows.

Danehy, Paul M.↗
NASA NTRS1973

Performance characteristics of a model VTOL lift fan in crossflow.

This paper presents a summary of principal results obtained from crossflow tests of a model 15-in.-diam lift fan installed in a wing in the NASA Lewis Research Center, 9 by 15 ft V/STOL Propulsion Wind Tunnel. Tests were run with and without exit louvers over a range of tunnel air speeds, fan speeds, and wing angle of attack. Fan thrust in crossflow was influenced by two principal factors: the effects of inflow distortion on blade-row performance, and changes in fan stage operating point brought about by changes in back pressure ratio. In this particular fan, flow separation on the inlet bellmouth did not appear to be a serious problem for crossflow operation.

Lieblein, S.↗
NASA NTRS1984

Computer studies of hybrid-slotted working sections with minimum interference at subsonic speeds

A series of computations on tunnel boundary-interference effects for hybrid-slotted working sections was performed using the WALINT code. The slots were modeled as lines of porosity with linear crossflow characteristics. The basic shape evaluated was for a rectangular section with height-to-width ratio = 0.835 and its companion in the duplex mode (half model testing) with height-to-width ratio = 0.6. A best overall basic configuration was determined with seven slots on each wall with open area ratio on each wall of 17.5%. For both full-span and half-model testing, the optimum solution required closing all but two slots on each of the half-walls parallel to the plane of the wing (equivalent to four slots on the full floor and ceiling). The results are presented here for the best configurations and are shown to be within the figure-of-merit range of + or - 0.04 in upwash, and + or - 0.1 in curvature for the Mach number range 0.6 to 0.85. Blockage effects are shown to be small.

Steinle, F. W., Jr.↗
NASA NTRS1988

Mass loss from S stars

The mass-loss process in S stars is studied using 65 S stars from the listing of Wing and Yorka (1977). The role of pulsations in the mass-loss process is examined. It is detected that stars with larger mass-loss rates have a greater amplitude of pulsations. The dust-to-gas ratio for the S stars is estimated as 0.002 and the average mass-loss rate is about 6 x 10 to the -8th solar masses/yr. Some of the properties of the S stars, such as scale height, surface density, and lifetime, are measured. It is determined that scale height is 200 pc; the total duration of the S star phase is greater than or equal to 30,000 yr; and the stars inject 3 x 10 to the -6th solar masses/sq kpc yr into the interstellar medium.

Jura, M.↗
NASA NTRS1999

Wind-Tunnel Overpressure Signatures From a Low-Boom HSCT Concept With Aft-Fuselage-Mounted Engines

A 1:300 scale wind-tunnel model of a conceptual High-Speed Civil Transport (HSCT) designed to generate a shaped, low-boom pressure signature on the ground was tested to obtain sonic-boom pressure signatures in the Langley Research Center Unitary Plan Wind Tunnel at a Mach number of 1.8 and a separation distance of about two body lengths or four wing-spans from the model. Two sets of engine nacelles representing two levels of engine technology were used on the model to determine the effects of increased nacelle volume. Pressure signatures were measured for (model lift)/(design lift) ratios of 0.5, 0.63, 0.75, and 1.0 so that the effect of lift on the pressure signature could be determined. The results of these tests were analyzed and used to discuss the agreement between experimental data and design expectations.

Mack, Robert J.↗
NASA NTRS2000

High Speed Vortex Flows

A review of the research conducted at the National Aeronautics and Space Administration (NASA), Langley Research Center (LaRC) into high-speed vortex flows during the 1970s, 1980s, and 1990s is presented. The data reviewed is for flat plates, cavities, bodies, missiles, wings, and aircraft. These data are presented and discussed relative to the design of future vehicles. Also presented is a brief historical review of the extensive body of high-speed vortex flow research from the 1940s to the present in order to provide perspective of the NASA LaRC's high-speed research results. Data are presented which show the types of vortex structures which occur at supersonic speeds and the impact of these flow structures to vehicle performance and control is discussed. The data presented shows the presence of both small- and large scale vortex structures for a variety of vehicles, from missiles to transports. For cavities, the data show very complex multiple vortex structures exist at all combinations of cavity depth to length ratios and Mach number. The data for missiles show the existence of very strong interference effects between body and/or fin vortices and the downstream fins. It was shown that these vortex flow interference effects could be both positive and negative. Data are shown which highlights the effect that leading-edge sweep, leading-edge bluntness, wing thickness, location of maximum thickness, and camber has on the aerodynamics of and flow over delta wings. The observed flow fields for delta wings (i.e. separation bubble, classical vortex, vortex with shock, etc.) are discussed in the context of' aircraft design. And data have been shown that indicate that aerodynamic performance improvements are available by considering vortex flows as a primary design feature. Finally a discussing of a design approach for wings which utilize vortex flows for improved aerodynamic performance at supersonic speed is presented.

Wood, Richard M.↗
NASA NTRS1982

Nacelle Aerodynamic and Inertial Loads (NAIL) project

Inlet airloads were measured by integration of pressures recorded at 252 locations on the right hand inboard (JT9D-7A) nacelle of the Boeing 747 aircraft. Pressures were recorded at 45 locations on the right hand outboard nacelle for comparison. Inertial loads were measured on both nacelles using accelerometers and rate gyros. Flight conditions included takeoffs at several gross weights, high g turns, and a simulated acceptance flight. Airloads at takeoff rotation were found to be larger than at any other normal service condition because of the combined effects of high angle of attack and high engine airflow. Inertial loads were smaller than previous estimates had indicated. A procedure is given for estimating inlet airloads at low speeds and high angles of attack for any underwing high bypass ratio turbofan installation approximately resembling the one tested. Flight procedure modifications are suggested that may result in better fuel economy retention in service. Pressures were recorded on the core cowls and pylons of both engine installations and on adjacent wing surfaces for use in development of computer codes for analysis of installed propulsion system aerodynamic drag interference effects.

Source record↗