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At least 577 records · Page 32

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NASA NTRS1983

Interaction of multiple supersonic jets with a transonic flow field

The influence of multiple high pressure, supersonic, radial or tangential jets, that are injected from the circumference of the base plane of an axisymmetric body, on its longitudinal aerodynamic coefficients in transonic flow is studied experimentally. The interaction of the jets with the body flow field increases the pressures on the forebody, thus altering its lift and static stability characteristics. It is shown that, within the range of parameters studied. This interaction has a stabilizing effect on the body. The contribution to lift and stability is significant at small angles of attack and decreases nonlinearly at higher angles when the crossflow mechanism becomes dominant.

Seginer, A.↗

NASA NTRS2002

Evolving Flow-Field Centers in Oblique Impacts

Ejection velocities and angles for vertical impacts are compared to different regions of 30 degree impact curtains using 3D PIV. Ejecta dynamics of the different regions require flow-field centers that change through space and time. Additional information is contained in the original extended abstract.

Anderson, J. L. B.↗

NASA NTRS1986

Computation of nozzle flow fields using the PARC2D Navier-Stokes code

Supersonic nozzles which operate at low Reynolds numbers and have large expansion ratios have very thick boundary layers at their exit. This leads to a very strong viscous/inviscid interaction upon the flow within the nozzle and the traditional nozzle design techniques which correct the inviscid core with a boundary layer displacement do not accurately predict the nozzle exit conditions. A full Navier-Stokes code (PARC2D) was used to compute the nozzle flow field. Grids were generated using the interactive grid generator code TBGG. All computations were made on the NASA MSFC CRAY X-MP computer. Comparison was made between the computations and in-house wall pressure measurements for CO2 flow through a conical nozzle having an area ratio of 40. Satisfactory agreement existed between the computations and measurements for a stagnation pressure of 29.4 psia and stagnation temperature of 1060 R. However, agreement did not exist at a stagnation pressure of 7.4 psia. Several reasons for the lack of agreement are possible. The computational code assumed a constant gas gamma whereas gamma for CO2 varied from 1.22 in the plenum chamber to 1.38 at the nozzle exit. Finally, it is possible that condensation occurred during the expansion at the lower stagnation pressure.

Collins, Frank G.↗

NASA NTRS1972

Surface pressure and inviscid flow field properties of the North American Rockwell delta-wing orbiter for nominal Mach number of 8, volume 2

The results are presented of a wind tunnel test program to determine surface pressures and flow field properties on the space shuttle orbiter configuration. The tests were conducted in September 1971. Data were obtained at a nominal Mach number of 8 and a free stream unit Reynolds number of 3.7 million per foot. Angle of attack was varied from 10 to 50 deg in 10-deg increments.

Matthews, R. K.↗

NASA NTRS1986

Pitot-Pressure Measurements in Flow Fields Behind a Rectangular Nozzle with Exhaust Jet for Free-Stream Mach Numbers of 0.00, 0.60, and 1.20

An investigation has been conducted in the Langley 16-Foot Transonic Tunnel to measure the flow field in and around the jet exhaust from a nonaxisymmetric nozzle configuration. The nozzle had a rectangular exit with a width-to-height ratio of 2.38. Pitot-pressure measurements were made at five longitudinal locations downstream of the nozzle exit. The maximum distance downstream of the exit was about 5 nozzle heights. These measurements were made at free-stream Mach numbers of 0.00, 0.60, and 1.20 with the nozzle operating at a ratio of nozzle total pressure to free-stream static pressure of 4.0. The jet exhaust was simulated with high-pressure air that had an exit total temperature essentially equal to the free-stream total temperature.

Putnam, L. E.↗

NASA NTRS1987

The effects of engine speed and injection characteristics on the flow field and fuel/air mixing in motored two-stroke diesel engines

A numerical analysis is presented on the effects of the engine speed, injection angle, droplet distribution function, and spray cone angle on the flow field, spray penetration and vaporization, and turbulence in a turbocharged motored two-stroke diesel engine. The results indicate that the spray penetration and vaporization, velocity, and turbulence kinetic energy increase with the intake swirl angle. Good spray penetration, vaporization, and mixing can be achieved by injecting droplets of diameters between 50 and 100 microns along a 120-deg cone at about 315 deg before top-dead-center for an intake swirl angle of 30 deg. The spray penetration and vaporization were found to be insensitive to the turbulence levels within the cylinder. The results have also indicated that squish is necessary in order to increase the fuel vaporization rate and mixing.

Nguyen, H. L.↗

NASA NTRS1987

Temperature measurement in a compressible flow field using laser-induced iodine fluorescence

The thermometric capability of a two-line fluorescence technique using iodine seed molecules in air is investigated analytically and verified experimentally in a known steady compressible flow field. Temperatures ranging from 165 to 295 K were measured in the flowfield using two iodine transitions accessed with a 30-GHz dye-laser scan near 543 nm. The effect of pressure broadening on temperature measurement is evaluated.

Fletcher, D. G.↗

NASA NTRS1998

Prediction of Particle Number Density and Particle Properties in the Flow Field Observed by the Nephelometer Experiment on the Galileo Probe

This report summarizes the work performed to assist in the analysis of data returned from the Galileo Probe's Nephelometer instrument. A computation of the flow field around the Galileo Probe during its descent through the Jovian atmosphere was simulated. The behavior of cloud particles that passed around the Galileo probe was then computed and the number density in the vicinity of the Nephelometer instrument was predicted. The results of our analysis support the finding that the number density of cloud particles was not the same in each of the four sampling volumes of the Nephelometer instrument. The number densities calculated in this study are currently being used to assist in the reanalysis of the data returned from the Galileo Probe.

Naughton, Jonathan W.↗

NASA NTRS1998

Prediction of Particle Number Density and Particle Properties in the Flow Field Observed by the Nephelometer Experiment on the Galileo Probe

Naughton, Jonathan W.↗

NASA NTRS1973

Heating and flow-field studies on a straight-wing hypersonic reentry vehicle at angles of attack from 20 to 80 deg with simulation of real-gas trends

Data are presented from a series of phase-change heat transfer and flow visualization tests at Mach 7.4, 8, and 10.3 in air, Mach 19.5 in nitrogen, Mach 20.3 in helium, and Mach 6 in tetrafluoromethane (CF4) on the windward surface of a straight wing hypersonic reentry configuration for angles of attack from 20 deg to 80 deg. The results indicate that: (1) for hypersonic stream Mach numbers, the flow field over the straight-wing configuration is essentially independent of Mach number, (2) transition Reynolds number decreases with increasing angle of attack, (3) at some critical angle of attack, the wing-shock standoff distance is greatly increased and the stagnation line moves downstream from the wing leading edge, (4) value of the critical angle of attack is very sensitive to the flow shock density ratio or effective gamma, and (5) at angles of attack above the critical value for all gases, the nondimensional level of heat transfer to the wing is higher for the higher shock density ratio flows.

Hunt, J. L.↗

NASA NTRS1971

An analysis of the flow field near the fuel injection location in a gas core reactor.

An analytical study is presented which shows the effects of large energy release and the concurrent high acceleration of inner stream fluid on the coaxial flow field in a gas core reactor. The governing equations include the assumptions of only radial radiative transport of energy represented as an energy diffusion term in the Euler equations. The method of integral relations is used to obtain the numerical solution. Results show that the rapidly accelerating, heat generating inner stream actually shrinks in radius as it expands axially.

Weinstein, H.↗

NASA NTRS1980

Calculation of two-dimensional inlet flow fields in a supersonic free stream by an implicit marching code with nonorthogonal mesh generation: User's manual

An implicit, shock-capturing finite-difference code which is used to calculate two-dimensional inlet flow fields in a supersonic free stream is explained. The Euler equations are subjected to general nonorthogonal transformation and a body-fitted coordinate system is employed. The mathematical formulation of the problem is given along with the numerical algorithm. Initial and boundary conditions, numerical stability, program limitations, and accuracy is discussed. An overall program logic as well as instructions for program use and operation are also furnished.

Biringen, S. H.↗

NASA NTRS1986

Experiments and modeling of dilution jet flow fields

Experimental and analytical results of the mixing of single, double, and opposed rows of jets with an isothermal or variable-temperature main stream in a straight duct are presented. This study was performed to investigate flow and geometric variations typical of the complex, three-dimensional flow field in the dilution zone of gas-turbine-engine combustion chambers. The principal results, shown experimentally and analytically, were the following: (1) variations in orifice size and spacing can have a significant effect on the temperature profiles; (2) similar distributions can be obtained, independent of orifice diameter, if momentum-flux ratio and orifice spacing are coupled; (3) a first-order approximation of the mixing of jets with a variable-temperature main stream can be obtained by superimposing the main-stream and jets-in-an-isothermal-crossflow profiles; (4) the penetration of jets issuing mixing is slower and is asymmetric with respect to the jet centerplanes, which shift laterally with increasing downstream distance; (5) double rows of jets give temperature distributions similar to those from a single row of equally spaced, equal-area circular holes; (6) for opposed rows of jets, with the orifice centerlines in line, the optimum ratio of orifice spacing to duct height is one-half the optimum value for single-side injection at the same momentum-flux ratiol and (7) for opposed rows of jets, with the orifice centerlines staggered, the optimum ratio of orifice spacing to duct height is twice the optimum value for single-side injection at the same momentum-flux ratio.

Holdeman, James D.↗

NASA NTRS1999

Investigation of the Rocket Induced Flow Field in a Rectangular Duct

Rocket-Based Combined Cycle (RBCC) concepts attempt to improve the performance of launch vehicles at all points in the launch trajectory and make highly reusable launch vehicles a reality. The Aerojet Strutjet RBCC concept consists of a variable geometry duct with internal, vertical struts that functions in ducted rocket, ramjet, scramjet, and pure rocket modes. These struts have rocket and turbine exhaust nozzles imbedded within them. The rocket flows create an ejector effect with the ingested air at subsonic flight velocities. In ramjet and scramjet modes, the fuel rich nozzle flows react with the ingested air producing an afterburner effect. Under a NASA Marshall Space Flight Center contract, the UAH Propulsion Research Center (PRC) has designed and built a Strutjet simulation facility. A scale model of a single strut has been built and is undergoing cold-flow testing to investigate the mixing of the rocket and turbine exhausts with the ingested air. A complementary experimental program is also underway to examine the induced flow-field generated by rocket nozzles confined in a rectangular duct. Characterizing the induced flow behavior is critical to understanding and optimizing the performance of future Strutjet-based RBCC propulsion systems. The proposed paper will present results from the rocket induced flow investigation.

Landrum, D. Brian↗

NASA NTRS1986

Mach 6 flow field surveys beneath the forebody of an airbreathing missile

Wall static, local stream static, and pitot pressure surveys were made on the windward side of a hypersonic airbreathing missile at full-scale length Reynolds numbers. In the inviscid part of the flow field, the experimental massflow ratios agreed with trends predicted by a three-dimensional method-of-characteristics solution. At a longitudinal station 3.5 diameters downstrea of the nose, the boundary layer was transitional or turbulent at zero incidence but became laminar as the angle of attack increased. The bell-shaped distribution of the boundary layer across the width of the body affected the mass flow distribution out to the bow shock and decreased the mass flow available the engine inlet.

Johnson, P. J.↗

NASA NTRS1982

Numerical solution of Space Shuttle Orbiter flow field

The supersonic, viscous laminar flow around the Space Shuttle Orbiter forebody has been computed with a parabolized Navier-Stokes code using a generalized coordinate transformation. The initial solution for the nose part of the Orbiter geometry was obtained with a three-dimensional time-dependent Navier-Stokes solver. It was necessary to employ a wind axis oriented coordinate system to obtain the initial solution with the time-dependent code. The generalized PNS technique used in this study allows the solution to be marched from the given initial data surface to any desired surface downstream. A grid point clustering scheme was employed to accurately describe the body shape by clustering points at the wing tip and at the wing body juncture. The computed heat transfer coefficients, pressure coefficients, and shock shapes are compared with the available experimental data for 0 degrees and 30 degrees angle of attack.

Venkatapathy, E.↗

NASA NTRS2008

Unified Application Vapor Screen Flow Visualization and Pressure Sensitive Paint Measurement Techniques to Vortex- and Shock Wave-Dominated Flow Fields

Laser vapor screen (LVS) flow visualization and pressure sensitive paint (PSP) techniques were applied in a unified approach to wind tunnel testing of slender wing and missile configurations dominated by vortex flows and shock waves at subsonic, transonic, and supersonic speeds. The off-surface cross-flow patterns using the LVS technique were combined with global PSP surface static pressure mappings to characterize the leading-edge vortices and shock waves that coexist and interact at high angles of attack (alpha). The synthesis of LVS and PSP techniques was also effective in identifying the significant effects of passive surface porosity and the presence of vertical tail surfaces on the flow topologies. An overview is given of LVS and PSP applications in selected experiments on small-scale models of generic slender wing and missile configurations in the NASA Langley Research Center (NASA LaRC) Unitary Plan Wind Tunnel (UPWT) and 8-Foot Transonic Pressure Tunnel (8-Foot TPT).

Erickson, Gary E.↗

NASA NTRS2010

Unified Application of Vapor Screen Flow Visualization and Pressure Sensitive Paint Measurement Techniques to Vortex- and Shock Wave-Dominated Flow Fields

Laser vapor screen (LVS) flow visualization and pressure sensitive paint (PSP) techniques were applied in a unified approach to wind tunnel testing of slender wing and missile configurations dominated by vortex flows and shock waves at subsonic, transonic, and supersonic speeds. The off-surface cross-flow patterns using the LVS technique were combined with global PSP surface static pressure mappings to characterize the leading-edge vortices and shock waves that coexist and interact at high angles of attack. The synthesis of LVS and PSP techniques was also effective in identifying the significant effects of passive surface porosity and the presence of vertical tail surfaces on the flow topologies. An overview is given of LVS and PSP applications in selected experiments on small-scale models of generic slender wing and missile configurations in the NASA Langley Research Center (NASA LaRC) Unitary Plan Wind Tunnel (UPWT) and 8-Foot Transonic Pressure Tunnel (8-Foot TPT).

Erickson, Gary E.↗