Search NASA⌕ Search

SEARCH · Search NASA

Results for “CONES”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS1987

Stability of axisymmetric boundary layers on sharp cones at hypersonic Mach numbers

The stability of sharp-cone boundary layers at zero angle of attack is investigated. Standard linear stability theory is used to perform a numerical study at an edge Mach number M(e) of 6.8 of normal-mode stability characteristics on an adiabatic-wall cone with special reference to the conditions of the stability experiment of Stetson et al. (1983). Comparisons of the calculations with experimental measurements bring out major areas of disagreement which remain to be resolved even in this simple case. Finally, a series of calculations of both cone and flat-plate N factors at M(e) of 4.5, 5.8, and 6.8 is used to study the relation between transition on a cone and flat plate based on stability theory.

Mack, Leslie M.↗
NASA NTRS1986

Hydroburst test of a carbon-carbon involute exit cone

A hydroburst test of the aft portion of the PAM-D exit cone and the test procedure are described in detail. The hydrostatic pressure required to buckle the cone was 9.75 psi. Meanwhile, the PAM-D exit cone was modeled using the finite element method and a theoretical bucking pressure (8.76 psi) was predicted using the SPAR finite element code. The modeling technique employed is discussed. By comparing the theoretical to predicted critical pressures, this report verifies the modeling technique and calculates a material knockdown factor for the carbon-carbon exit cone.

Sullivan, Roy M.↗
NASA NTRS1989

Semidiscrete Galerkin modelling of compressible viscous flow past a circular cone at incidence

A numerical study of the laminar and compressible boundary layer, about a circular cone in a supersonic free stream, is presented. It is thought that if accurate and efficient numerical schemes can be produced to solve the boundary layer equations, they can be joined to numerical codes that solve the inviscid outer flow. The combination of these numerical codes is competitive with the accurate, but computationally expensive, Navier-Stokes schemes. The primary goal is to develop a finite element method for the calculation of 3-D compressible laminar boundary layer about a yawed cone. The proposed method can, in principle, be extended to apply to the 3-D boundary layer of pointed bodies of arbitrary cross section. The 3-D boundary layer equations governing supersonic free stream flow about a cone are examined. The 3-D partial differential equations are reduced to 2-D integral equations by applying the Howarth, Mangler, Crocco transformations, a linear relation between viscosity, and a Blasius-type of similarity variable. This is equivalent to a Dorodnitsyn-type formulation. The reduced equations are independent of density and curvature effects, and resemble the weak form of the 2-D incompressible boundary layer equations in Cartesian coordinates. In addition the coordinate normal to the wall has been stretched, which reduces the gradients across the layer and provides high resolution near the surface. Utilizing the parabolic nature of the boundary layer equations, a finite element method is applied to the Dorodnitsyn formulation. The formulation is presented in a Petrov-Galerkin finite element form and discretized across the layer using linear interpolation functions. The finite element discretization yields a system of ordinary differential equations in the circumferential direction. The circumferential derivatives are solved by an implicit and noniterative finite difference marching scheme. Solutions are presented for a 15 deg half angle cone at angles of attack of 5 and 10 deg. The numerical solutions assume a laminar boundary layer with free stream Mach number of 7. Results include circumferential distribution of skin friction and surface heat transfer, and cross flow velocity distributions across the layer.

Meade, Andrew James, Jr.↗
NASA NTRS1991

Navier-Stokes predictions of pitch damping for axisymmetric shell using steady coning motion

Previous theoretical investigations have proposed that the side force and moment acting on a body of revolution in steady coning motion could be related to the pitch-damping force and moment. In the current research effort, this approach is applied to produce predictions of the pitch damping for axisymmetric shell. The flow fields about these projectiles undergoing steady coning motion are successfully computed using a parabolized Navier-Stokes computational approach which makes use of a rotating coordinate frame. The governing equations are modified to include the centrifugal and Coriolis force terms due to the rotating coordinate frame. From the computed flow field, the side moments due to coning motion, spinning motion, and combined spinning and coning motion are used to determine the pitch-damping coefficients. Computations are performed for two generic shell configurations, a secant-ogive-cylinder and a secant-ogive-cylinder-boattail.

Weinacht, Paul↗
NASA NTRS1992

Spectral solution of inviscid supersonic flows over wedges and axisymmetric cones

A shock-fitted multidomain spectral collocation method is used to solve both steady and unsteady inviscid supersonic flows over bodies. New aspects of the method include two subdomain interface types and a zonal solution procedure to get efficient convergence to steady-state in supersonic regions. Steady-state examples include flow over a sharp cone, a hyperbolic cone and a hyperbolic wedge. For the cone, the exact flow solution is used to show that the method is spectrally accurate. As an example of an unsteady flow, a calculation of the interaction of a free-stream hot ring with the flow over a sharp cone is presented.

Kopriva, David A.↗
NASA NTRS1992

Plume effects on the flow around a blunted cone at hypersonic speeds

Tests at M = 8.2 show that a simulated rocket plume at the base of a blunted cone can cause large areas of separated flow, with dramatic effects on the heat transfer rate distribution. The plume was simulated by solid discs of varying sizes or by an annular jet of gas. Flow over the cone without a plume is fully laminar and attached. Using a large disc, the boundary layer is laminar at separation at the test Reynolds number. Transition occurs along the separated shear layer and the boundary layer quickly becomes turbulent. The reduction in heat transfer associated with a laminar separated region is followed by rising values as transition occurs and the heat transfer rates towards the rear of the cone substantially exceed the values obtained without a plume. With the annular jet or a small disc, separation occurs much further aft, so that heat transfer rates at the front of the cone are comparable with those found without a plume. Downstream of separation the shear layer now remains laminar and the heat transfer rates to the surface are significantly lower than the attached flow values.

Atcliffe, P.↗
NASA NTRS1958

Investigation of an underslung half-cone inlet with compression surface mounted outboard from fuselage at Mach numbers of 1.5, 1.8, and 2.0

An investigation was conducted to determine the performance of an underslung half-cone inlet mounted on a missile forebody model with the compression surface outboard from the fuselage. The inlet was designed for shock-on-lip operation at Mach number 2.0 with 25 degree half-angle spike. The cowling was attached to the fuselage through the boundary-layer plow and served as part of the fuselage boundary-layer diverter system. The performance of the half-cone inlet was compared with that of a scoop-type inlet and a normal-wedge inlet on a maximum-thrust-minus-drag basis. The increase in pressure recovery obtained with the half-cone inlet over that obtained with the reference inlets offset the slightly higher drags observed over the Mach number range for the half-cone so that the performance of this configuration was equal to that of the other inlets at Mach number 2.0 and was slightly superior at the lower Mach numbers. For a particular configuration, a peak pressure recovery of 0.879 was obtained at Mach number 2.0, zero angle of attack, and 4-percent throat bleed; the subcritical stability was 16 percent. Use of a fuselage-mounted boundary-layer splitter plate ahead of the inlet did not improve the stability. Subcritical distortion values were below 10 percent for all configurations. (author)

Yeager, Richard A↗
NASA NTRS1972

Nonlinear Aerodynamics of Bodies in Coning Motion

A numerical method for computing the nonlinear inviscid flowfield surrounding a body performing coning motion is described. The method permits accurate computation of the aerodynamic moment due to one of the four motions characterizing an arbitrary nonplanar motion. Results of computations for a slender circular cone in coning motion are presented, and show good agreement with experiment for angles of attack up to twice the cone half-angle. The computational results display significant departure of the side moment from the linear theory value with increasing angle of attack, but agree well with experimental measurements. This indicates that the initial nonlinear behavior of the aerodynamic moment is determined primarily by the inviscid flow.

Schiff, Lewis B.↗
NASA NTRS1959

Stability Investigation of a Blunted Cone and a Blunted Ogive with a Flared Cylinder Afterbody at Mach Numbers from 0.30 to 2.85

A cone with a blunt nose tip and a 10.7 deg cone half angle and an ogive with a blunt nose tip and a 20 deg flared cylinder afterbody have been tested in free flight over a Mach number range of 0.30 to 2.85 and a Reynolds number range of 1 x 10(exp 6) to 23 x 10(exp 6). Time histories, cross plots of force and moment coefficients, and plots of the longitudinal force,coefficient, rolling velocity, aerodynamic center, normal- force-curve slope, and dynamic stability are presented. With the center-of-gravity location at about 50 percent of the model length, the models were both statically and dynamically stable throughout the Mach number range. For the cone, the average aerodynamic center moved slightly forward with decreasing speeds and the normal-force-curve slope was fairly constant throughout the speed range. For the ogive, the average aerodynamic center remained practically constant and the normal-force-curve slope remained practically constant to a Mach number of approximately 1.6 where a rising trend is noted. Maximum drag coefficient for the cone, with reference to the base area, was approximately 0.6, and for the ogive, with reference to the area of the cylindrical portion, was approximately 2.1.

Coltrane, Lucille C.↗
NASA NTRS1961

Experimental Determination of the Recovery Factor and Analytical Solution of the Conical Flow Field for a 20 deg Included Angle Cone at Mach Numbers of 4.6 and 6.0 and Stagnation Temperatures to 2600 degree R

The local recovery factor was determined experimentally along the surface of a thin-walled 20 deg included angle cone for Mach numbers near 6.0 at stagnation temperatures between 1200 deg R and 2600 deg R. In addition, a similar cone configuration was tested at Mach numbers near 4.5 at stagnation temperatures of approximately 612 deg R. The local Reynolds number based on flow properties at the edge of the boundary layer ranged between 0.1 x 10(exp 4) and 3.5 x 10(exp 4) for tests at temperatures above 1200 deg R and between 6 x 10(exp 4) and 25 x 10(exp 4) for tests at temperatures near 612 deg R. The results indicated, generally, that the recovery factor can be predicted satisfactorily using the square root of the Prandtl number. No conclusion could be made as to the necessity of evaluating the Prandtl number at a reference temperature given by an empirical equation, as opposed to evaluating the Prandtl number at the wall temperature or static temperature of the gas at the cone surface. For the tests at temperatures above 1200 deg R (indicated herein as the tests conducted in the slip-flow region), two definite trends in the recovery data were observed - one of increasing recovery factor with decreasing stagnation pressure, which was associated with slip-flow effects and one of decreasing recovery factor with increasing temperature. The true cause of the latter trend could not be ascertained, but it was shown that this trend was not appreciably altered by the sources of error of the magnitude considered herein. The real-gas equations of state were used to determine accurately the local stream properties at the outer edge of the boundary layer of the cone. Included in the report, therefore, is a general solution for the conical flow of a real gas using the Beattie-Bridgeman equation of state. The largest effect of temperature was seen to be in the terms which were dependent upon the internal energy of the gas. The pressure and hence the pressure drag terms were unaffected.

Pfyl, Frank A.↗
NASA NTRS2002

Origins of Small Volcanic Cones on Mars

Studies of volcanic cones identified in the MGS data indicate a range of possible origins, from primary vent constructs (cinder cones, tuff cones) to rootless cones formed by lava-ice interaction. Additional information is contained in the original extended abstract.

Fagents, S. A.↗
NASA NTRS2003

Icelandic Pseudocraters as Analogs to some Volcanic Cones on Mars

Pseudocraters are rootless vents formed by the interaction of lava flows with surface or near-surface water. This interaction can produce mild explosions and the accumulation of scoria and spatter into small constructs. Pseudocraters in several localities in Iceland were examined in the field and compared to similar appearing features observed on Mars. The Icelandic pseudocrater cones in this study range in size from 6 to 70 m in diameter, have summit craters which range from 2 to 28 m in diameter (many cones lack craters entirely), and have flanks that am either concave- up or convex-up. The size and spacing of Icelandic pseudo-craters might be a function of the availability of water, in which larger, closely spaced features result from efficient lava-water interaction, as suggested by the environments in which the features formed. Possible Martian pseudocrater cones in Amamnis Planitia range in diameter from 30 to 180 m and have craters 12 to 80 m in diameter. A numerical model for volcanic explosions was adapted to study the formation of pseudocraters under terrestrial and Martian conditions. The results suggest that explosions forming Martian cones require significantly less water (calculated masses am less by a factor of 4 to 16) than those forming Icelandic pseudokers, despite their larger sizes, This is attributed to the low gravity and atmospheric pressure in the Mars environment and is consistent with the likely lower abundance of water, which might be present as interstitial ice at shallow depths in the regolith. Locations of potential pseudocraters on Mars at latitudes as low as approximately 8 degrees N, imply the presence of crustal ice stores at the time of their formation.

Greeley, Ronald↗
NASA NTRS2007

Receptivity of Supersonic Boundary Layers Due To Acoustic Disturbances Over Blunt Cones

Receptivity and stability of supersonic boundary layers over a 5-degree straight cone with a blunt tip are numerically investigated at a free stream Mach number of 3.5 and at a high Reynolds number of 106/inch. Both the steady and unsteady solutions are obtained by solving the full Navier-Stokes equations using the 5th-order accurate weighted essentially non-oscillatory (WENO) scheme for space discretization and using third-order total-variation-diminishing (TVD) Runge-Kutta scheme for time integration. The linear stability results showed that bluntness has less stabilizing effects on the stability of boundary layers over cones than on flat plates and wedges. The unsteady simulations of the interaction of plane threedimensional acoustic waves with the cone showed that the modulation of wavelength and the generation of instability waves first occurred near the leading edge in the plane where the constant acoustic phase lines are perpendicular to the cone axis. Further downstream, this instability region spreads in the azimuthal direction from this plane.

Balakumar, P.↗
NASA NTRS2009

Stability of Supersonic Boundary Layers on a Cone at an Angle of Attack

The stability and receptivity of three-dimensional supersonic boundary layers over a 7deg sharp tipped straight cone at an angle of attack of 4.2deg is numerically investigated at a free stream Mach number of 3.5 and at two high Reynolds numbers, 0.25 and 0.50x10(exp 6)/inch. The generation and evolution of stationary crossflow vortices are also investigated by performing simulations with three-dimensional roughness elements located on the surface of the cone. The flow fields with and without the roughness elements are obtained by solving the full Navier-Stokes equations in cylindrical coordinates using the fifth-order accurate weighted essentially non-oscillatory (WENO) scheme for spatial discretization and using the third-order total-variation-diminishing (TVD) Runge-Kutta scheme for temporal integration. Stability computations reveal that the azimuthal wavenumbers are in the range of m approx. 25-50 for the most amplified traveling disturbances and in the range of m approx. 40-70 for the stationary disturbances. The N-Factor computations predicted that transition would occur further forward in the middle of the cone compared to the transition fronts near the windward and the leeward planes. The simulations revealed that the crossflow vortices originating from the nose region propagate towards the leeward plane. No perturbations were observed in the lower part of the cone.

Balakumar, Ponnampalam↗
NASA NTRS2010

Analysis of Instabilities in Non-Axisymmetric Hypersonic Boundary Layers Over Cones

Hypersonic flows over circular cones constitute one of the most important generic configurations for fundamental aerodynamic and aerothermodynamic studies. In this paper, numerical computations are carried out for Mach 6 flows over a 7-degree half-angle cone with two different flow incidence angles and a compression cone with a large concave curvature. Instability wave and transition-related flow physics are investigated using a series of advanced stability methods ranging from conventional linear stability theory (LST) and a higher-fidelity linear and nonlinear parabolized stability equations (PSE), to the 2D eigenvalue analysis based on partial differential equations. Computed N factor distribution pertinent to various instability mechanisms over the cone surface provides initial assessments of possible transition fronts and a guide to corresponding disturbance characteristics such as frequency and azimuthal wave numbers. It is also shown that strong secondary instability that eventually leads to transition to turbulence can be simulated very efficiently using a combination of advanced stability methods described above.

Li, Fei↗
NASA NTRS2010

Stability of Hypersonic Boundary Layers on a Cone at an Angle of Attack

The stability and receptivity of a three-dimensional hypersonic boundary layer over a 7deg half-angle straight cone at an angle of attack of 6deg is numerically investigated at a freestream Mach number of 6.0 and a Reynolds number of 10.4x10(exp 6)/m. The generation and evolution of stationary crossflow vortices are also investigated by performing simulations with three-dimensional roughness elements located on the surface of the cone. The flow fields with and without the roughness elements are obtained by solving the full Navier- Stokes equations in cylindrical coordinates using a fifth-order accurate weighted essentially non-oscillatory (WENO) scheme for spatial discretization and a third-order total-variation-diminishing (TVD) Runge-Kutta scheme for temporal integration. Stability computations produced azimuthal wavenumbers in the range of m approx. 20-50 for the most amplified traveling disturbances and in the range of m approx.30-70 for the stationary disturbances. The frequency of the unstable second-mode ranges from 400 kHz to 900 kHz along the windward ray. The N-Factor computations predicted transition would occur more forward on the sides of the cone as compared to the transition fronts near the windward and the leeward rays. The simulations also show the crossflow vortices originating from the nose region propagate towards the leeward ray. No perturbations were observed toward the windward half of the cone.

Balakumar, P.↗
NASA NTRS2016

Measurements in a Transitioning Cone Boundary Layer at Freestream Mach 3.5

An experimental study was conducted in the Supersonic Low-Disturbance Tunnel to investigate naturally-occurring instabilities in a supersonic boundary layer on a 7 deg half- angle cone. All tests were conducted with a nominal freestream Mach number of M(sub infinity) = 3:5, total temperature of T(sub 0) = 299:8 K, and unit Reynolds numbers of Re(sub infinity) x 10(exp -6) = 9:89, 13.85, 21.77, and 25.73 m(exp -1). Instability measurements were acquired under noisy- ow and quiet- ow conditions. Measurements were made to document the freestream and the boundary-layer edge environment, to document the cone baseline flow, and to establish the stability characteristics of the transitioning flow. Pitot pressure and hot-wire boundary- layer measurements were obtained using a model-integrated traverse system. All hot- wire results were single-point measurements and were acquired with a sensor calibrated to mass ux. For the noisy-flow conditions, excellent agreement for the growth rates and mode shapes was achieved between the measured results and linear stability theory (LST). The corresponding N factor at transition from LST is N 3:9. The stability measurements for the quiet-flow conditions were limited to the aft end of the cone. The most unstable first-mode instabilities as predicted by LST were successfully measured, but this unstable first mode was not the dominant instability measured in the boundary layer. Instead, the dominant instabilities were found to be the less-amplified, low-frequency disturbances predicted by linear stability theory, and these instabilities grew according to linear theory. These low-frequency unstable disturbances were initiated by freestream acoustic disturbances through a receptivity process that is believed to occur near the branch I locations of the cone. Under quiet-flow conditions, the boundary layer remained laminar up to the last measurement station for the largest Re1, implying a transition N factor of N greater than 8:5.

King, Rudolph A.↗
NASA NTRS2021

Boundary Layer Instabilities Over a Cone-Cylinder-Flare Model at Mach 6

Computations are performed to investigate the boundary-layer instabilities over a sharp cone-cylinder-flare model at zero degrees angle of attack. The model geometry and the flow conditions are selected to match the experiments conducted in the Boeing/AFOSR Mach 6 Quiet Tunnel (BAM6QT) at Purdue University. The geometry consists of a nominally sharp 5-degree half-angle cone, followed by a cylindrical segment and then a 10-degree flare. An axisymmetric separation bubble is generated as a result of the laminar shock/boundary-layer interaction in the cylinder-flare region. The comparison of the laminar flow solution and the schlieren images shows a remarkable agreement between the respective locations of both the boundarylayer edge and the reattachment shock. The predicted heat flux distribution is also in agreement with the measured values downstream of the reattachment location. The analysis of convective and global instabilities is performed for flare half angles equal to 8, 10, and 12 degrees and nosetip radii equal to 0.1, 1, and 5 mm. The linear amplification of first and second Mack mode instabilities that begin to amplify in the cone region are computed with a combination of the parabolized stability equations (PSE) and the harmonic linearized Navier-Stokes equations (HLNSE). The predicted frequency spectra of the surface pressure fluctuations associated with both planar and oblique instability waves are compared with the measured spectra at the various locations of the PCB and Kulite sensors. The comparison shows that the computational analysis captures the distinct lobes within the disturbance amplification spectra measured in the experiments, but some differences in amplification characteristics are noted at low frequencies. Overall, the oblique disturbances are found to be more amplified than the planar disturbances. To our knowledge, this represents the first successful comparison between convective instability analysis and measured surface pressure fluctuations for a hypersonic configuration with a separation bubble. Finally, the global instability analysis shows that the laminar flow becomes supercritical for flare half angles larger than 8 degrees. The unstable global mode for the experimental configuration of a 10 degrees flare and a sharp nosetip cone corresponds to a stationary three-dimensional disturbance that is concentrated in the recirculation region and achieves its maximum growth rate for an azimuthal wavenumber of 5.

Boundary layer transition↗