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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 55 records · Page 3
All sourcesNASA NTRSDOE OSTI
NASA NTRS1975

The effect of Reynolds number on boattail drag

An investigation has been conducted in the Langley pilot transonic cryogenic tunnel to determine the effects of varying Reynolds number on boattail drag at subsonic speeds. Six boattailed cone-cylinder nacelle models were tested with the jet exhaust simulated by a cylindrical sting. Reynolds number was varied from about 2.6 million to 132 million by changing model length and unit Reynolds number. Boattail pressure coefficient distributions show that increasing Reynolds number tends to make the pressure coefficients in the expansion region more negative and the pressure coefficients in the recompression region more positive. These two effects were compensating and as a result there was little or no effect of Reynolds number on the pressure drag of the isolated boattails.

Reubush, D. E.↗
NASA NTRS1975

The effect of Reynolds number on the boattail drag of two wing-body configurations

An investigation has been conducted in the Langley 1/3-meter transonic cryogenic tunnel to determine the effects of varying Reynolds number on the boattail drag of wing-body configurations at subsonic speeds. Two boattailed cone-cylinder nacelle models were tested with a 60 deg delta wing at an angle of attack of 0 deg. Reynolds number, based on model length, was varied from about 2.5 million to 67 million. Even though the presence of the wing had large effects on the boattail pressure coefficients, the results of this investigation were similar to those previously found for a series of isolated boattails. Boattail pressure coefficients in the expansion region became more negative with increasing Reynolds number, while those in the recompression region became more positive. These two effects were compensating, and as a result, there was virtually no effect of Reynolds number on boattail pressure drag.

Reubush, D. E.↗
NASA NTRS1976

An experimental and analytical investigation of effect on isolated boattail drag of varying Reynolds numbers up to 130,000,000

An investigation was conducted to determine whether large Reynolds number effects occur on isolated boattails, including an analytical study and tests in a 1/3-meter transonic cryogenic tunnel. This investigation was conducted at an angle of attack of 0 deg at Mach numbers from 0.6 to 0.9 for Reynolds numbers up to 130 million. Results indicate that as the Reynolds number was increased, the static pressure coefficients in the expansion region of the boattail became more negative whereas those in the recompression region became more positive. These two trends were compensating and, as a result, there was only a small effect (if any) of Reynolds numbers on boattail pressure drag.

Reubush, D. E.↗
NASA NTRS1976

Effect of Reynolds number on the subsonic boattail drag of several wing-body configurations

An investigation was conducted in a transonic cryogenic tunnel to determine the effect of varying Reynolds number on the boattail drag of several wing-body configurations. This study was made at 0 deg angle of attack at Mach numbers from 0.6 to 0.9 for Reynolds numbers up to 67 x 1 million (based on distance from the nose to the start of the boattail). Results indicate that as the Reynolds number was increased the boattail static pressure coefficients in the expansion region of the boattail became more negative while those in the recompression region became more positive. Results show that there was only a small effect of Reynolds number of boattail pressure drag.

Reubush, D. E.↗
NASA NTRS1976

Laminar supersonic flow over a backstep - A numerical solution at higher Reynolds numbers

The Allen-Cheng solution of the flow over a backward facing step is extended to Reynolds numbers up to 16,000 and to inflow boundary-layer height ratios as low as 0.1 by moving the downstream boundary into the recompression region and by smoothing the resulting errors. The boundary conditions in the supersonic outer flow and the downstream boundary conditions in the wake are determined by an extrapolation procedure. Computational results are compared with relevant experimental data. Fair agreement is found between the calculated base pressures and the experimental values, whereas agreement between heat transfer rates appears to be qualitative only.

Kronzon, Y.↗
NASA NTRS1976

Numerical simulation of supersonic and hypersonic turbulent compression corner flows using relaxation models

Relaxation turbulence eddy viscosity models are incorporated to solve the complete Navier-Stokes equations for supersonic and hypersonic flows over a two-dimensional compression corner. The system of equations is solved by a time-split, second-order accurate numerical scheme. Details of relaxation process are studied, and several relaxation models are tested and compared. Good improvement in the prediction of upstream pressure propagation is obtained for a Mach number of 2.96, and Reynolds number of 10,000,000, with wedge angle of 25 deg. However, the application of the relaxation models to hypersonic flow at Mach number 8.66 and Reynolds number of 22,000,000, with highly cooled wall, shows unfavorable effects on heat transfer and skin friction in the reattachment and recompression regions.

Hung, C. M.↗
NASA NTRS1976

Viscid-inviscid interaction of incompressible separated flows

A flow model has been devised to deal with the viscid-inviscid interaction of a class of two-dimensional incompressible separated flow problems. It is suggested that the corresponding inviscid flow of these problems is described by the free streamline theory with few unspecified parameters and their values are, in turn, determined by the viscous flow considerations. The problem of a flow past a backward facing step is selected for study in detail. The viscous flow components of turbulent jet mixing, recompression, and reattachment are delineated and studied individually. When they are later combined, it is found that the point of reattachment behaves as a saddle-point-type singularity in the system of differential equations describing the viscous flow process. This feature is employed to the determination of the aforementioned free parameters and thus the establishment of the overall corresponding inviscid flow field. The resulting base pressure coefficient for the specific case agrees reasonably well with the available experimental data.

Chow, W. L.↗
NASA NTRS1979

Fluid dynamic problems associated with air-breathing propulsive systems

A brief account of research activities on problems related to air-breathing propulsion is made in this final report for the step funded research grant NASA NGL 14-005-140. Problems include the aircraft ejector-nozzle propulsive system, nonconstant pressure jet mixing process, recompression and reattachment of turbulent free shear layer, supersonic turbulent base pressure, low speed separated flows, transonic boattail flow with and without small angle of attack, transonic base pressures, Mach reflection of shocks, and numerical solution of potential equation through hodograph transformation.

Chow, W. L.↗
NASA NTRS1979

Numerical solution for perpendicular sonic hydrogen injection into a ducted supersonic airstream

This note discusses a computer program being developed to study the flow field near opposing perpendicular fuel injectors in scramjets. The MacCormack time-split, finite difference relaxation technique was used to solve the full two-dimensional compressible Navier-Stokes equations along with energy and species equations. By using this technique, a program was developed to consider the turbulent nonreacting flow of hydrogen and air in a rectangular duct. A damping term, proportional to the second derivative of pressure and temperature, was used to produce a stable solution behind the hydrogen jet in the neighborhood of the recompression shock. A case using actual conditions encountered in current scramjet design was analyzed, with results agreeing qualitatively with experimental observations.

Drummond, J. P.↗
NASA NTRS1981

Effects of a modified leading edge on noise and boundary-layer transition in a rod-wall sound shield at Mach 5

A version of a rod wall sound shield was tested in the Mach 5 pilot quiet tunnel over a range of unit Reynolds numbers from 10 to 35 million per meter. The model was modified by inclining the leading edge plates to produce an initial 2 deg expansion to ascertain the sensitivity of boundary layer transition to leading edge disturbances. Rod surface pitot pressures, mean free stream pitot pressures, and static pressures on the rods and plenum walls were measured. Hot-wire measurements were also made in the model and nozzle free stream at a unit Reynolds number of 15 million per meter. The surface pitot pressures indicated that transition was much farther forward than for the previous tests due to the leading edge modification and minor fabrication flaws in the model. Early boundary layer transition on the rods was confirmed by hot-wire measurements which showed much higher noise levels in the free stream shield flow when compared with results from previous tests. Mean pitot pressure surveys within the shielded region inside the model indicated that there was an overexpansion and recompression that would limit the streamwise length of undisturbed flow to about 13 cm along the centerline.

Creel, T. R., Jr.↗
NASA NTRS1969

Aerodynamic Characteristics of an 11-Percent-Thick Symmetrical Supercritical Airfoil at Mach Numbers Between 0.30 and 0.85

The aerodynamic characteristics of an 11 percent thick symmetrical supercritical airfoil were determined in an 8 foot transonic pressure tunnel over a Mach number range of 0.30 to 0.85. The Reynolds number of the tests, based on the airfoil chord, varied with Mach number over a range of 360 million to 774 million. The geometric angle of attack varied from -0.5 to 10.5 deg. The abrupt drag rise for the supercritical airfoil at zero normal force conditions occurs at a Mach number just above 0.82. The corresponding drag rise Mach number for a conventional NACA 0012 airfoil is approximately 0.70. At zero normal force conditions, the level of supervelocity over the supercritical airfoil is considerably reduced from that for the NACA 0012 airfoil. Also, the shock wave for the supercritical airfoil is substantially weaker than that for the NACA 0012 airfoil. For a Mach number of 0.82 and zero normal force, the flow over the present airfoil is supercritical; however, there is no discernible shock wave in the flow, indicating near isentropic recompression.

Blackwell, J. A., Jr.↗
NASA NTRS1982

Optical studies of shock generated transient supersonic base flows

A shock tube employing interferometric and schlieren techniques is used to study transient base flow phenomena following shock wave passage over two plane bluff bodies: a hemicircular cylinder and a cylinder with the Galileo Jovian probe profile. An attempt is made to understand the physics of transition from transient to steady state flow, and to provide code verification for a study employing the Illiac IV computer. Transient base flow interactions include a series of shock diffraction, regular, and Mach reflections, coupled with boundary layer development, separation, and recompression. Vorticity generation and transport underlie these features. The quantitative verification of the computer code includes comparisons of transient pressure and density fields, near wake geometries, and bow shock standoff distances.

Liang, P.-Y.↗
NASA NTRS1982

Evolution of shock-induced pressure on a flat-face/flat-base body and afterbody flow separation

The time-dependent, compressible Reynolds-averaged, Navier-Stokes equations are applied to solve an axisymmetric supersonic flow around a flat-face/flat-base body with and without a sting support. Important transient phenomena, not yet well understood, are investigated, and the significance of the present solution to the phenomena is discussed. The phenomena, described in detail, are as follows: the transient formation of the bow and recompression shock waves; the evolution of a pressure buildup due to diffraction of the incident shock wave in the forebody and afterbody regions, including the luminosity accompanying the pressure buildup; the separation of the flow as influenced by pressure buildup; the location of the separation and the reattachment points; and the transient period of the shock-induced base flow. The important influence of the nonsteady (transient) and steady flow on the aerodynamic characteristics, radiative heat transfer, and, thus, on the survivability or safeguard problems for an aircraft fuselage, missile, or planetary entry probe at very high flight speeds is described.

Yoshikawa, K. K.↗
NASA NTRS1983

SC3 - A wing concept for supersonic maneuvering

A supersonic wing design concept that overcomes the problems associated with nonlinear inviscid pressure distribution is presented together with a summary of the supporting theoretical and experimental development program. The key design feature is a conical panel supporting a controlled supercritical conical camber (SC3) crossflow expansion and recompression which permits lift on the upper surface to be obtained without producing an adverse pressure gradient or crossflow shock wave strong enough to separate the boundary layer. The role of aerodynamics in the design implementation is discussed, as are the concept development, the initial design validation on a fabricated and tested SC3 design, and the body and canard interaction effect. The design and testing of an isolated wing with a planform and thickness distribution representative of a practical application are described. Wind-tunnel results show that significant performance gains can be obtained for fighter aircraft using this concept.

Mason, W. H.↗
NASA NTRS1983

A new method to measure homogeneous nucleation rates in shock tubes

The centered expansion wave of a shock tube is utilized to expand and supersaturate a condensable vapor in small concentration in an inert carrier gas. The supersaturated state, located at the rear of the expansion wave, is preserved for a controlled period and then terminated by a recompressing shock wave. During the period of supersaturation, condensation nuclei are formed homogeneously. The nucleation rate is measured as a function of supersaturation by a Mie-light scattering technique. The method is tested using water and the results are compared with classical nucleation theory.

Peters, F.↗
NASA NTRS1988

Structure of a reattaching supersonic shear flow

A Mach 1.83 fully developed turbulent boundary layer with boundary layer thickness, free stream velocity, and Reynolds number of 7.5 mm, 476 m/s, and 6.2 x 10 to the 7th/m, respectively, was separated at a 25.4-mm backward step and formed a shear layer. Fast-response pressure transducers, schlieren photography, and LDV were used to study the structure of this reattaching shear flow. The preliminary results show that large-scale relatively organized structures with limited spanwise extent form in the free shear layer. Some of these structures appear to survive the recompression and reattachment processes, while others break down into smaller scales and the flow becomes increasingly three-dimensional. The survived large-scale structures lose their organization through recompression/reattachment, but regain it after reattachment. The structures after reattachment form a 40-45-degree angle relative to the free stream and deteriorate gradually as they move downstream.

Samimy, M.↗
NASA NTRS1993

Numerical analysis of base flowfield at high altitude for a four-engine clustered nozzle configuration

The objective of this study is to benchmark a four-engine clustered nozzle base flowfield with a computational fluid dynamics (CFD) model. The CFD model is a pressure based, viscous flow formulation. An adaptive upwind scheme is employed for the spatial discretization. The upwind scheme is based on second and fourth order central differencing with adaptive artificial dissipation. Qualitative base flow features such as the reverse jet, wall jet, recompression shock, and plume-plume impingement have been captured. The computed quantitative flow properties such as the radial base pressure distribution, model centerline Mach number and static pressure variation, and base pressure characteristic curve agreed reasonably well with those of the measurement. Parametric study on the effect of grid resolution, turbulence model, inlet boundary condition and difference scheme on convective terms has been performed. The results showed that grid resolution and turbulence model are two primary factors that influence the accuracy of the base flowfield prediction.

Wang, Ten-See↗
NASA NTRS1946

Application of the method of characteristics to supersonic rotational flow

A system for calculating the physical properties of supersonic rotational flow with axial symmetry and supersonic rotational flow in a two-dimensional field was determined by use of the characteristics method. The system was applied to the study of external and internal flow for supersonic inlets with axial symmetry. For a circular conical inlet the shock that occurred at the lip of the inlet became stronger as it approached the axis of the inlet and became a normal shock at the axis. The region in which strong shock occurred increased with increase of the angle of internal cone at the lip of the inlet. For an inlet with a central body the method of characteristics was applied to the design of an internal channel shape that, theoretically, results in very efficient recompression in the inlet. It was shown that if an effuser is connected with the diffuser a body of revolution with very small shock-wave drag can be determined. (author)

Ferri, Antonio↗