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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 253 records · Page 14
All sourcesNASA NTRSDOE OSTI
NASA NTRS1983

Comparison of measured and predicted flight effects on high-bypass coaxial jet exhaust noise

A semi-empirical model for predicting the noise generated by conventional-velocity profile coaxial jets is compared with full scale flight data and model scale simulated flight data for high bypass nozzles. The prediction model was shown to agree with small scale static data for primary jet velocities from 215 to 795 m/s for a wide range of area, temperature, and velocity ratios between streams. However, there were insufficient model nozzle, simulated flight data available at that time to permit validation of the flight effects prediction. The comparisons presented demonstrate that the prediction method is also valid in flight.

Stone, J. R.↗
NASA NTRS1983

Comparison of measured and predicted flight effects on high-bypass coaxial jet exhaust noise

A recently developed semi-empirical model for predicting the noise generated by conventional-velocity-profile coaxial jets is compared with full-scale flight data and model-scale simulated flight data for high-bypass nozzles. The prediction model has been shown to agree with small-scale static data for primary jet velocities from 215 to 795 m/s for a wide range of area, temperature and velocity ratios between streams. However, there were insufficient model nozzle, simulated flight data available at that time to permit validation of the flight effects prediction. The comparisons presented in this paper demonstrate that the prediction method is also valid in flight.

Stone, J. R.↗
NASA NTRS1987

Static internal performance of a two-dimensional convergent-divergent nozzle with thrust vectoring

A parametric investigation of the static internal performance of multifunction two-dimensional convergent-divergent nozzles has been made in the static test facility of the Langley 16-Foot Transonic Tunnel. All nozzles had a constant throat area and aspect ratio. The effects of upper and lower flap angles, divergent flap length, throat approach angle, sidewall containment, and throat geometry were determined. All nozzles were tested at a thrust vector angle that varied from 5.60 tp 23.00 deg. The nozzle pressure ratio was varied up to 10 for all configurations.

Bare, E. Ann↗
NASA NTRS1974

Noise and static performance characteristics of a STOL aircraft jet flap

Static noise and performance tests were conducted on a 1/4-scale jet flap model with a multilobe nozzle of array area ratio of 2.7. The model nozzle and flap tested were a two-dimensional section of a distributed blowing system similar to previously investigated augmentor wing systems without the upper shroud and intake door. Noise data were measured with the nozzle alone and also during attached flow conditions with the flap at two turning angles representing takeoff and approach conditions. The noise data are scaled to a 200,000-lb TOGW four-engine airplane and are presented in terms of perceived noise level and one-third octave band sound pressure level. Comparisons are made with the noise levels produced by an augmentor wing airplane fitted with a three-element acoustically lined augmentor flap. The static performance is presented in terms of thrust recovery and effective turning angle.

Harkonen, D. L.↗
NASA NTRS1984

Flow in sigmoid diffusers of moderate curvature

Developing laminar and turbulent flows have been measured in two diffusers with S- and C-shaped center-lines, small divergence angle and moderate curvature. The square inlet cross-section expanded in the plane of curvature to a rectangle with an exit-to-inlet area ratio of 1.5. Laser-Doppler anemometry was used to measure three velocity components and, in turbulent flow, the associated fluctuations and cross-correlations. Wall static pressure measurements are also reported. Flow visualization did not reveal the presence of any recirculation regions. Pressure-driven secondary flows are present in both ducts, but are generally smaller than those measured in similar ducts of uniform cross-section area.

Rojas, J.↗
NASA NTRS1994

Dynamics and statics of nonaxisymmetric liquid bridges

We finished the construction of the experimental apparatus and the design and testing of some of the visualization and data acquisition techniques. Experimental work focused on three areas: force measurements, loss of stability to nonaxisymmetric bridges, and vibration behavior. The experimental work is summarized in section 2. Selected results from our force measurement experiments are outlined in section 3. In addition we worked on the theory of the dynamic stability of axisymmetric bridges and undertook numerical simulation of the effects of inclined gravity vectors on the minimum volume stability limit for static bridges. The results and status of our theoretical work and numerical simulation are described in section 4. Papers published and in preparation, conference presentations, etc., are described in section 5. Work planned for the third year is discussed in section 6. References cited in the report are listed in section 7.

Alexander, J. Iwan D.↗
NASA NTRS2003

Exploratory Calibration of Adjustable-Protrusion Surface-Obstacle (APSO) Skin Friction Vector Gage

The design of an adjustable-protrusion surface-obstacle (APSO) skin friction vector gage is presented. Results from exploratory calibrations conducted in laminar and turbulent boundary layers at the Washington University Low-Speed Wind Tunnel and for turbulent boundary layers at speeds up to Mach 2 on the ceiling of the NASA Glenn Research Center 8- X 6-ft Supersonic Wind Tunnel are also discussed. The adjustable-height gage was designed to yield both the magnitude and direction of the surface shear stress vector and to measure the local static pressure distribution. Results from the NASA test show good correlation for subsonic and low supersonic conditions covering several orders of magnitude in terms of the adopted similarity variables. Recommendations for future work in this area consist of identifying the physical parameters responsible for the disagreement between the university and NASA data sets, developing a compressibility correction specific to the APSO geometry, and examining the effect that static pressure distribution and skewed boundary layers have on the results from the APSO.

Hakkinen, Raimo J.↗
NASA NTRS1974

A parametric study of effect of forebody shape on flow angularity at Mach 8

Flow angularity and static pressure measurements have been made on the lower surface of nine forebody models that simulate the bottom forward surface of a hypersonic aircraft. Measurements were made in an area of the forebody that represents the location of an inlet of a scramjet engine. A parametric variation of the forebody surface investigated the effect of: (1) spanwise curvature; (2) longitudinal curvature; and (3) planform shape on both flow angularity and static pressure distribution. Results of each of the three parametric variations of geometry were compared to those for the same flat delta forebody. Spanwise curvature results showed that a concave shape and the flat delta had the lowest flow angularity and lowest rate of increase in flow angularity with angle of attack. Longitudinal curvature results showed a convex surface to give the better flow at the higher angles of attack. The better of the two planform shapes tested was a convex elliptical shape. Limited flow field calculations were made at angles of attack using a three dimensional, method-of-characteristics program. In general, at all angles of attack there was agreement between data and theory.

Johnson, C. B.↗
NASA NTRS1975

A surface-wave investigation of the rupture mechanism of the Gobi-Altai (4 December 1957) earthquake

Long period records of multiple Love waves from the 1957 earthquake in Mongolia at Pasadena are analyzed and compared to synthetic seismograms, generated by the method of Kanamori. A fit in the time domain shows that the records are not consistent with the previous solution, achieved through a frequency domain analysis of directivity. The solution asks for a shorter rupture of 270 km at a velocity of 3.5 km/s. The focal parameters are constrained by updating all the reported first motion and are found to be: Strike = 103 deg, Dip = 53 deg, Slip = 32 deg. A seismic moment of 1.8 10 to the 28th power dynes-cm is obtained. These figures are also consistent with a time domain analysis of Love waves at Palisades and Strasbourg, and of Rayleigh waves at Pasadena, with a directivity study of Love waves at Pasadena, and with static deformation and isoseismal data. A discussion is given of the relation between moment, magnitude and rupture area, and a comparison is made with other events in the same region: It is concluded that this earthquake does not exhibit an intra-plate behavior, but rather compares better with inter-plate events, such as the great Assam earthquake.

Okal, E. A.↗
NASA NTRS1981

Evaluation of superplastic forming and co-diffusion bonding of Ti-6Al-4V titanium alloy expanded sandwich structures

The application of the superplastic forming/diffusion bonding (SPF/DB) process to supersonic cruise research is investigated. The capability of an SPF/DB titanium structure to meet the structural requirements of the inner wing area of the NASA arrow-wing advanced supersonic transport design is evaluated. Selection of structural concepts and their optimization for minimum weight, SPF/DB process optimization, fabrication of representative specimens, and specimen testing and evaluation are described. The structural area used includes both upper and lower wing panels, where the upper wing panel is used for static compression strength evaluation and the lower panel, in tension, is used for fracture mechanics evaluations. The individual test specimens, cut from six large panels, consist of 39 static specimens, 10 fracture mechanics specimens, and one each full size panel for compression stability and fracture mechanics testing. Tests are performed at temperatures of -54 C (-65 F), room temperature, and 260 C (500 F).

Arvin, G. H.↗
NASA NTRS1983

Passive shock wave/boundary layer control for transonic airfoil drag reduction

An investigation of the passive shock wave/boundary layer interaction control for reducing the drag in 12 percent thick circular arc and 14 percent thick supercritical airfoils was conducted in a 3 in. x 15.4 in. transonic wind tunnel at transonic Mach numbers. A porous surface with a cavity beneath it was positioned on the area of the airfoils, mounted on the test section bottom wall, where the shock wave occurs. The static pressure distributions over the airfoil, the wake impact pressure data for determining the profile drag, and the Schlieren photographs for porous surface airfoils are presented and compared with the results for solid surface airfoils. With the porous surface the normal shock wave for solid surface was changed to a lambda shock wave system, and the wake impact pressure data indicated an appreciable drag reduction for both airfoils with the porous surface at transonic speeds while causing little or no loss of lift. The effect of porosity and cavity size is investigated and off-design performance is discussed.

Bahi, L.↗
NASA NTRS1984

Porosity effect on supercritical airfoil drag reduction by shock wave/boundary layer control

An investigation of the passive shock wave/boundary layer control for reducing the drag of 14 percent-thick supercritical airfoil was conducted in the 3 in. x 15.4 in. RPI Transonic Wind Tunnel at transonic Mach numbers. Various porous surfaces with a cavity beneath it was positioned on the area of the airfoil, mounted on the test section bottom wall, where the shock wave occurs. The static pressure distributions over the airfoil, the wake impact pressure survey for determining the profile drag and the Schlieren photographs for porous surfaces are presented and compared with the results for solid surface airfoil. With a uniform porosity surface the normal shock wave for solid surface was changed to a lambda shock wave, and the wake impact pressure data indicated an appreciable drag reduction at transonic Mach numbers. For a free stream Mach number of 0.81 the profile drag coefficient for the airfoil top surface with uniform porosity was 46 percent lower than for the solid surface airfoil.

Nagamatsu, H. T.↗
NASA NTRS1991

Preliminary calibration of a generic scramjet combustor

The results of a preliminary investigation of the combustion of hydrogen fuel at hypersonic flow conditions are provided. The tests were performed in a generic, constant-area combustor model with test gas supplied by a free-piston-driven reflected-shock tunnel. Static pressure measurements along the combustor wall indicated that burning did occur for combustor inlet conditions of P(static) approximately equal to 19kPa, T(static) approximately equal to 1080 K, and U approximately equal to 3630 m/s with a fuel equivalence ratio approximately equal to 0.9. These inlet conditions were obtained by operating the tunnel with stagnation enthalpy approximately equal to 8.1 MJ/kg, stagnation pressure approximately equal to 52 MPa, and a contoured nozzle with a nominal exit Mach number of 5.5.

Jacobs, P. A.↗
NASA NTRS1992

Numerical Modeling of a Vortex Stabilized Arcjet

Arcjet thrusters are being actively considered for use in Earth orbit maneuvering applications. Experimental studies are currently the chief means of determining an optimal thruster configuration. Earlier numerical studies have failed to include all of the effects found in typical arcjets including complex geometries, viscosity, and swirling flow. Arcjet geometries are large area ratio converging nozzles with centerbodies in the subsonic portion of the nozzle. The nozzle walls serve as the anode while the centerbody functions as the cathode. Viscous effects are important because the Reynolds number, based on the throat radius, is typically less than 1,000. Experimental studies have shown that a swirl or circumferential velocity component stabilizes a constricted arc. This dissertation describes the equations governing flow through a constricted arcjet thruster. An assumption that the flowfield is in local thermodynamic equilibrium leads to a single fluid plasma temperature model. An order of magnitude analysis reveals the governing fluid mechanics equations are uncoupled from the electromagnetic field equations. A numerical method is developed to solve the governing fluid mechanics equations, the Thin Layer Navier-Stokes equations. A coordinate transformation is employed in deriving the governing equations to simplify the application of boundary conditions in complex geometries. An axisymmetric formulation is employed to include the swirl velocity component as well as the axial and radial velocity components. The numerical method is an implicit finite-volume technique and allows for large time steps to reach a converged steady-state solution. The inviscid fluxes are flux-split, and Gauss-Seidel line relaxation is used to accelerate convergence. Converging-diverging nozzles with exit-to-throat area ratios up to 100:1 and annular nozzles were examined. Quantities examined included Mach number and static wall pressure distributions, and oblique shock structures. As the level of swirl and viscosity in the flowfield increased the mass flow rate and thrust decreased. The technique was used to predict the flow through a typical arcjet thruster geometry. Results indicate swirl and viscosity play an important role in the complex geometry of an arcjet.

Pawlas, Gary E.↗
NASA NTRS2014

Wind-US Results for the AIAA 2nd Propulsion Aerodynamics Workshop

This presentation contains Wind-US results presented at the 2nd Propulsion Aerodynamics Workshop. The workshop was organized by the American Institute of Aeronautics and Astronautics, Air Breathing Propulsion Systems Integration Technical Committee with the purpose of assessing the accuracy of computational fluid dynamics for air breathing propulsion applications. Attendees included representatives from government, industry, academia, and commercial software companies. Participants were encouraged to explore and discuss all aspects of the simulation process including the effects of mesh type and refinement, solver numerical schemes, and turbulence modeling. The first set of challenge cases involved computing the thrust and discharge coefficients for a 25deg conical nozzle for a range of nozzle pressure ratios between 1.4 and 7.0. Participants were also asked to simulate two cases in which the 25deg conical nozzle was bifurcated by a solid plate, resulting in vortex shedding (NPR=1.6) and shifted plume shock (NPR=4.0). A second set of nozzle cases involved computing the discharge and thrust coefficients for a convergent dual stream nozzle for a range of subsonic nozzle pressure ratios. The workshop committee also compared the plume mixing of these cases across various codes and models. The final test case was a serpentine inlet diffuser with an outlet to inlet area ratio of 1.52 and an offset of 1.34 times the inlet diameter. Boundary layer profiles, wall static pressure, and total pressure at downstream rake locations were examined.

Wind-US↗
NASA NTRS1978

Effects of inlet airframe integration on the inlet of an upper surface blowing four-engine STOL aircraft

Inlet and nacelle static pressures were measured on a 0.55-scale model of the Quiet Short-Haul Research Airplane (QSRA) in the Ames Research Center's 40- by 80-Foot Wind Tunnel. This model is powered by four JT-15D engines located above the wing with closely spaced adjacent inlets. A fifth JT-15D engine in the fuselage provides boundary-layer control air. Each inlet was instrumented with four to eight rows of axial pressure taps located between X/R approximately plus or minus 1. The tests simulated a broad range of aircraft operating conditions, including engine-out, with lift coefficients from 0.8 to 10.0. Results indicate that the inlets perform well under most operating conditions with little interaction between inlets when the aircraft is moving. Potential problem areas identified are high sideslip angle during approach and an interaction effect between adjacent inlets with high mass flows in static conditions.

Shovlin, M. D.↗
NASA NTRS1971

Static longitudinal aerodynamic characteristics of close-coupled wing-canard configurations at Mach numbers from 1.60 to 2.86

An experimental investigation was made in the Mach number range from 1.60 to 2.86 to determine the static longitudinal aerodynamic characteristics of close-coupled wing-canard configurations. Three canards, ranging in exposed planform area from 17.5 to 30.0 percent of the wing reference area, were employed in this investigation. The canards were either located in the plane of the wing or in a position 18.5 percent of the wing mean geometric chord above the wing plane. Most data obtained were for a model with a 60 deg leading-edge-sweep wing; however, a small amount of data were obtained for a 44 deg leading-edge-sweep wing. The model utilized two balances to isolate interference effects between wing and canard. In general, it was determined that at angle of attack for all configurations investigated with the canard in the plane of the wing an unfavorable interference exists which causes the additional lift on the canard generated by a canard deflection to be lost on the wing due to an increased downwash at the wing from the canard. Further, this interference decreased somewhat with increasing Mach number. Raising the canard above the plane of the wing also greatly decreased the interference of the canard deflection on the wing lift. However, at Mach 2.86 the presence of the canard in the high position had a greater unfavorable interference effect at high angles of attack than the canard in the wing plane. This interference resulted in the in-plane canard having better trimmed performance at Mach 2.86 for the same center-of-gravity location.

Dollyhigh, S. M.↗
NASA NTRS1973

Mechanisms for Mars dust storms.

Characteristics of the Mars global dust storm are reviewed. At the Mariner 9 encounter, the dust consisted of highly absorbing particles distributed rather uniformly up to great height (about 50 km). These observations together with temperature distributions inferred from the Mariner 9 IRIS by Hanel et al. (1972) are used to estimate global wind systems during the dust storm. The global distribution and direction of light surface streaks indicate that the axisymmetric circulation was a dominant part of flow during the dust storm. The axisymmetric winds may become strong enough to raise dust over wide areas of Mars' tropics under unusual conditions: the incoming solar radiation must be near its seasonal maximum, the static stability must be low, and the atmosphere must be able to absorb and re-emit a sizeable fraction of the incoming radiation. Strong winds around the periphery of the retreating south polar cap would be driven by the temperature gradient at the cap edge and by the mass outflow due to subliming CO2. These polar winds could generate local dust storms, raising the general level of dustiness, and providing the conditions necessary for onset of a global dust storm.

Leovy, C. B.↗