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Results for “SLENDER CONE”

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 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1963

Effects of Mass Addition on the Stability of Slender Cones at Hypersonic Speeds

The effect of mass addition on the flow over bodies moving at hypersonic speeds has been studied by several investigators (e.g., Cresci and Libby). In most of this work, primary attention logically has been directed toward the effects of foreign-gas injection on heat transfer and pressure distributions, and, principally for this reason, most of the work ha been done at zero angle of attack. The foreign gas can be provided either by some active injection system or by the action of an ablation heat shield. With increasing rates of injection, the basic flow about the body can be affected significantly. One such effect was observed in the paper by Cresci and Libby, where it was shown that the shockwave standoff distance can be increased by gas injection at the nose of a body.

SLENDER BODY↗
NASA NTRS1990

Laser transit anemometer measurements on a slender cone in the Langley unitary plan wind tunnel

A laser transit anemometer (LTA) system was used to probe the boundary layer on a slender (5 degree half angle) cone model in the Langley unitary plan wind tunnel. The anemometer system utilized a pair of laser beams with a diameter of 40 micrometers spaced 1230 micrometers apart to measure the transit times of ensembles of seeding particles using a cross-correlation technique. From these measurements, boundary layer profiles around the model were constructed and compared with CFD calculations. The measured boundary layer profiles representing the boundary layer velocity normalized to the edge velocity as a function of height above the model surface were collected with the model at zero angle of attack for four different flow conditions, and were collected in a vertical plane that bisected the model's longitudinal center line at a location 635 mm from the tip of the forebody cone. The results indicate an excellent ability of the LTA system to make velocity measurements deep into the boundary layer. However, because of disturbances in the flow field caused by onboard seeding, premature transition occurred implying that upstream seeding is mandatory if model flow field integrity is to be maintained. A description and results of the flow field surveys are presented.

Humphreys, William M., Jr.↗
NASA NTRS1974

Fluctuating surface pressure characteristics on slender cones in subsonic, supersonic, and hypersonic Mach-number flow

A program is presented to determine the flow-induced unsteady aerodynamic loads on space-shuttle orbiter configurations in the early reentry phase. Experiments on 8 degree half-angle cone/flat-base configurations were conducted at free-stream Mach numbers of 6 and 15 and Reynolds numbers corresponding to shuttle reentry. Data are available at free-stream Mach-numbers of 0.67, 2.5, 3, 4, 6, 15, and 22. It is shown that empirical expressions could be developed for the Mach-number dependence of the overall fluctuation pressure under an attached turbulent boundary layer and under a separated base-flow region.

Heller, H. H.↗
NASA NTRS1977

Flight boundary layer transition measurements on a slender cone at Mach 20

Experimental flight boundary layer transition data have been obtained on a 3.962-m-long, 5 deg half-angle cone with an initial nose radius of 0.254 cm. The data were obtained during re-entry from altitudes of approximately 30.480 to 18.288 km at a free-stream Mach number of 20. The free-stream Reynolds number varied from 6.56 x 10 to the 6th/m to 52.5 x 10 to the 6th/m, and the total enthalpy from about 18.3 to 16.9 MJ/kg. The locations of the beginning and end of transition were determined by the intersection of curves faired throug the laminar, transitional, and turbulent heating-rate data. The temperature-history technique for determining transition as currently used (sharp break in curve) was shown to compare unfavorably with the heating-rate-distribution method. The heating-rate-history technique, which is proportional to the temperature derivative and consequently more sensitive to perturbations, gives better agreement with the heating-rate distribution transition results.

Wright, R. L.↗
NASA NTRS1981

Flight experiments with a slender cone at angle of attack

The three-dimensional leeward separation about a 5 deg semi-angle cone at an 11 deg angle of attack was investigated in flight, in the wind tunnel, and by numerical computations. The test conditions were Mach numbers of 0.6, 1.5, and 1.8 at Reynolds numbers between 7 and 10 million based on free-stream conditions and a 30-inch wetted length or surface. The surface conditions measured included mean static and fluctuating pressures; skin friction magnitudes and separation line positions were obtained using obstacle blocks. The mean static pressures from flight and wind tunnel were in good agreement. The computed results gave the same distributions, but were slightly more positive in magnitude. The experimentally measured primary and secondary separation line locations compared closely with computed results. There were substantial differences in level and in trend between the surface root-mean-square pressure fluctuations obtained in flight and in the wind tunnel, due, it is thought, to a relatively high acoustic disturbance level in the tunnel compared with the quiescent conditions in flight.

Peake, D. J.↗
NASA NTRS1982

Flight, Wind Tunnel, and Numerical Experiments with a Slender Cone at Incidence

The three-dimensional leeward separation about a 5-deg semi-angle cone of 11 deg angle of incidence was Investigated in night, in the wind tunnel, and by numerical computations. The test conditions were Mach numbers of 0.6, 1.5, and 1.8 at Reynolds numbers between 7 and 10 million based on freestream conditions and a 76.2-cm (30-in.) length of surface. The surface pressure conditions measured included those of fluctuating and mean static, as well as recovery pressures generated by obstacle blocks to provide skin friction and separation-line locations. The mean static pressures from flight and wind tunnel were in reasonably good agreement. The computed results gave the same distributions, but were slightly more positive in magnitude. The experimentally measured primary and secondary separation line locations compared closely with computed results. There were substantial differences In level between the surface root-mean-square pressure fluctuations obtained in night and in the wind tunnel, due, It Is thought, to a relatively high acoustic disturbance level in the tunnel compared with the quiescent atmospheric conditions in night.

Peake, David J.↗
NASA NTRS2002

Boundary Layer Transition on Slender Cones in Conventional and Low Disturbance Mach 6 Wind Tunnels

An experimental investigation was conducted on a 5-degree half-angle cone and a 5-degree half-angle flared cone in a conventional Mach 6 wind tunnel to examine the effects of facility noise on boundary layer transition. The influence of tunnel noise was inferred by comparing transition onset locations determined from the present test to that previously obtained in a Mach 6 low disturbance quiet tunnel. Together, the two sets of experiments are believed to represent the first direct comparison of transition onset between a conventional and a low disturbance wind tunnel using a common test model and transition detection technique. In the present conventional hypersonic tunnel experiment, separate measurements of heat transfer and adiabatic wall temperatures were obtained on the conical models at small angles of attack over a range of Reynolds numbers, which resulted in laminar, transitional, and turbulent flow. Smooth model turbulent heating distributions are compared to that obtained with transition forced via discrete surface roughness. The model nosetip radius was varied to examine the effects of bluntness on transition onset. Despite wall to total temperature differences between the transient heating measurements and the adiabatic wall temperature measurement, the two methods for determining sharp cone transition onset generally yielded equivalent locations. In the 'noisy' mode of the hypersonic low disturbance tunnel, transition onset occurred earlier than that measured in the conventional hypersonic tunnel, suggesting higher levels of freestream acoustic radiation relative to the conventional tunnel. At comparable freestream conditions, the transition onset Reynolds number under low disturbance conditions was a factor of 1.3 greater than that measured on flared cone in the LaRC conventional hypersonic tunnel and a factor of 1.6 greater that the flared cone run in the low disturbance tunnel run 'noisy'. Navier-Stokes mean flow computations and linear stability analysis were conducted to assess the experimental results and have indicated N factors associated with sharp flared cone transition onset to be approximately a factor of 2 lower than that inferred from the corresponding low disturbance tunnel measurements.

Horvath, Thomas J.↗
NASA NTRS1951

Aerodynamic Characteristics of a Slender Cone-cylinder Body of Revolution at a Mach Number of 3.85

An experimental investigation of the aerodynamics of a slender cone-cylinder body of revolution was conducted at a Mach number of 3.85 for angles of attack of 0 degree to 10 degrees and a Reynolds number of 3.85x10(exp 6). Boundary-layer measurements at zero angle of attack are compared with the compressible-flow formulations for predicting laminar boundary-layer characteristics. Comparison of experimental pressure and force values with theoretical values showed relatively good agreement for small angles of attack. The measured mean skin-friction coefficients agreed well with theoretical values obtained for laminar flow over cones.

FLOW, LAMINAR↗
NASA NTRS1993

Symmetry breaking in vortical flows over cones - Theory and numerical experiments

A stability analysis suggests that inviscid incompressible flow, independent from angle of attack and regardless whether attached or separated, over slender cones is only marginally stable in regions of decelerating circumferential flow. Reducing slenderness or surface curvature lowers the energy level of harmonic perturbations, and, thus, reduces their impact on the overall stability of flows over slender cones. Associating the notion of instabilities in such flows with the onset of vortex asymmetries provides a model for explaining a variety of flow phenomena in Navier-Stokes simulations of laminar incompressible flows over three right circular cones at moderate to high angles of attack.

Hartwich, Peter M.↗
NASA NTRS1994

Symmetry breaking in vortical flows over cones -- theory and numerical experiments

A stability analysis suggests that inviscid incompressible flow, independent from angle of attack and regardless whether attached or separated, over slender cones is only marginally stable in regions of decelerating circumferential flow. Reducing slenderness or surface curvature lowers the frequency spectrum of the harmonic perturbations and, thus, reduces their impact on the overall stability of flows over slender cones. Associating the notion of instabilities in such flows with the onset of vortex asymmetries provides a model for explaining a variety of flow phenomena in Navier-Stokes simulations of laminar incompressible flows over three right circular cones at moderate to high angles of attack.

Hartwich, Peter M.↗
NASA NTRS1992

Effects of nose bluntness and angle of attack on slender bodies in hypersonic flows

The effects of angle of attack and nose bluntness on the flow field and wall quantities are investigated for hypersonic flows of air over slender bodies. The bodies considered are slender cones and straight biconic configurations. The numerical procedures used are based on the solution of complete Navier-Stokes equations in the nose region and parabolized Navier-Stokes equations in the downstream region. Results are obtained for a wide range of free stream conditions in which the gas behind the shock is treated as perfect. The flow field variables and surface quantities show significant differences when the angle of attack and nose bluntness are varied. The postshock flow field is studied in detail from the contour plots of Mach number, density, and temperature. Flow separation is observed on the leeward plane for an on-axis, 12.84 deg/7 deg (fore-cone and aft-cone angles) biconic geometry at 12 deg angle of attack. Also, the windward and leeward heating rates for the fore-cone section decrease by a factor of four and five, respectively, when the nose bluntness is increased by an order of magnitude. The effect of nose bluntness for slender cone persists as far as 200 nose radii downstream.

Tiwari, S. N.↗