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At least 163 records · Page 9

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NASA NTRS1991

The CONE program - An overview

The Cryogenic Orbital Nitrogen Experiment (CONE) is a liquid nitrogen cryogenic storage and supply system demonstration placed in orbit by the National Space Transportation System (NSTS) Orbiter and operated as an in-bay payload whose objective is to demonstrate needed critical components and technologies. A conceptual approach has been developed and an overview of the CONE program is described which includes the following: (1) a definition of the background and scope of the technology objectives being investigated, (2) a description of the payload design and operation, major features and rationale for the experiments being conducted, and (3) the justification for CONE relating to potential near-term benefits and risk mitigation for future systems.

Bailey, William J.↗

NASA NTRS1957

A discussion of cone and flat-plate Reynolds numbers for equal ratios of the laminar shear to the shear caused by small velocity fluctuations in a laminar boundary layer

By use of the linear theory of boundary-layer stability and Schlichting's formula for the maximum amplification of a disturbance, an approximate relation is derived between the Reynolds number on a cone and the Reynolds number on a flat plate for equal closeness to transition. The indication is that the ratio of the cone Reynolds number for transition, based on the distance to the cone apex, to the plate Reynolds number for transition, based on the distance to the leading edge, is not in general equal to 3, as has been suggested by other investigators, but varies from 3 when transition occurs at the minimum critical Reynolds number to unity when transition occurs at a large multiple of the critical Reynolds number.

FLOW, TURBULENT↗

NASA NTRS1955

Transonic Flow Past Cone Cylinders

Experimental results are presented for transonic flow post cone-cylinder, axially symmetric bodies. The drag coefficient and surface Mach number are studied as the free-stream Mach number is varied and, wherever possible, the experimental results are compared with theoretical predictions. Interferometric results for several typical flow configurations are shown and an example of shock-free supersonic-to-subsonic compression is experimentally demonstrated. The theoretical problem of transonic flow past finite cones is discussed briefly and an approximate solution of the axially symmetric transonic equations, valid for a semi-infinite cone, is presented.

Solomon, George E↗

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 NTRS1961

Laminar Heat-Transfer and Pressure Measurements at a Mach Number of 6 on Sharp and Blunt 15 deg Half-Angle Cones at Angles of Attack Up to 90 deg

Two circular conical configurations having 15 deg half-angles were tested in laminar boundary layer at a Mach number of 6 and angles of attack up to 90 deg. One cone had a sharp nose and a fineness ratio of 1.87 and the other had a spherically blunted nose with a bluntness ratio of 0.1428 and a fineness ratio of 1.66. Pressure measurements and schlieren pictures of the flow showed that near-conical flow existed up to an angle of attack of approximately 60 deg. At angles of attack above 70 deg high-pressure areas were present near the base and the bow shock wave was considerably curved. Comparison of the results with simply applied theories showed that on the stagnation line pressures may be predicted by Newtonian theory, and heat transfer by local yawed-cylinder theory based on the yaw angle of the windward generator and the local radius of the cone. Base effects increased the heat transfer in a region extending forward approximately 15 to 30 percent of the windward generator. Circumferential pressure distributions were higher than the corresponding Newtonian distribution and a better prediction was obtained by modifying the theory to match the pressure at 90 deg from the windward generator to that on the surface of the cone at an angle of attack of 0 deg. Circumferential heat-transfer distributions were predicted satisfactorily up to about 60 deg from the stagnation line by using Lees' heat-flux distribution based on the Newtonian pressure. The effects of nose bluntness at large angles of attack were very small in the region beyond two nose radii from the point of tangency.

Conti, Raul J.↗

NASA NTRS1961

Longitudinal Force and Moment Data at Mach Numbers from 0.60 to 1.40 for a Family of Elliptic Cones with Various Semiapex Angles

An investigation has been made to determine the aerodynamic characteristics of four elliptic cones having plan-form semiapex angles ranging from about 9 to 31 deg., and also for one of these cones modified on the upper surface to reduce the base area by about one half. The tests were made for angles of attack from about -2 to +21 deg., at Mach numbers from 0.60 to 1.40, and for a constant Reynolds number of 1.4 million, based on the length of the models. For each model, lift, pitching-moment, and drag coefficients, and lift-drag ratios are presented for the forebody, and axial-force coefficients are presented for the base. Calculated lift and pitching- moment curves for the elliptic cones, and lift-curve slopes for each model at supersonic Mach numbers are shown for comparison with the corresponding experimental values. Lift-drag ratios are also given for the forebody and base combined. These data are presented without discussion.

Stivers, Louis S., Jr.↗

NASA NTRS1961

Measurement and Empirical Correlation of Transpiration-Cooling Parameters on a 25 degree Cone in a Turbulent Boundary Layer in Both Free Flight and a Hot-Gas Jet

Transpiration-cooling parameters are presented for a turbulent boundary layer on a cone configuration with a total angle of 250 which was tested in both free flight and in an ethylene-heated high-temperature jet at a Mach number of 2.0. The flight-tested cone was flown to a maximum Mach number of 4.08 and the jet tests were conducted at stagnation temperatures ranging from 937 R to 1,850 R. In general, the experimental heat transfer was in good agreement with the theoretical values. Inclusion of the ratio of local stream temperature to wall temperature in the nondimensional flow rate parameter enabled good correlation of both sets of transpiration data. The measured pressure at the forward station coincided with the theoretical pressure over a sharp cone; however, the measured pressure increased with distance from the nose tip.

Walton, Thomas E., Jr.↗

NASA NTRS2002

Mars Global Surveyor Data Analysis Program. Origins of Small Volcanic Cones: Eruption Mechanisms and Implications for Water on Mars

The goal of the proposed work was to determine the origins of small volcanic cones observed in Mars Global Surveyor (MGS) data, and their implications for regolith ice stores and magma volatile contents. For this 1-year study, our approach involved a combination of: Quantitative morphologic analysis and interpretation of Mars Orbiter Camera (MOC) and Mars Orbiter Laser Altimeter (MOLA) data; Numerical modeling of eruption processes responsible for producing the observed features; Fieldwork on terrestrial analogs in Iceland. Following this approach, this study succeeded in furthering our understanding of (i) the spatial and temporal distribution of near-surface water ice, as defined by the distribution and sizes of rootless volcanic cones ("pseudocraters"), and (ii) the properties, eruption conditions, and volatile contents of magmas producing primary vent cones.

Fagents, Sarah A.↗

NASA NTRS1947

Cones in Supersonic Flow

In the case of cones in axially symmetric flow of supersonic velocity, adiabatic compression takes place between shock wave and surface of the cone. Interpolation curves betwen shock polars and the surface are therefore necessary for the complete understanding of this type of flow. They are given in the present report by graphical-numerical integration of the differential equation for all cone angles and airspeeds.

Hantzsche, W.↗

NASA NTRS1961

Effect of Nose Bluntness on Transition for a Cone and a Hollow Cylinder at Mach Numbers 1.41 and 2.01

An investigation has been made to determine the effects of nose bluntness on boundary-layer transition for a cone with an included angle of 10 degrees and for a hollow cylinder. The tests were conducted at Mach numbers of 1.41 and 2.01 for free-stream Reynolds numbers per foot ranging from 1 x 10(exp 6) to 9 x 10(exp 6). The investigation was made with the use of schlieren photography for which the models were aligned with the free stream. For the 10 degree cone, the favorable effects of nose blunting were so small at both test Mach numbers as to be lost within the experimental accuracy. For small amounts of nose blunting on the hollow cylinder, for which the ratio of bluntness height to transition distance for the sharp-leading-edge cylinder was relatively small, there was little, if any, effect of blunting on transition. For somewhat larger values of this ratio, nose blunting had a favorable effect on transition. The magnitude of the favorable effect was dependent upon the size and the shape of the bluntness, and the maximum increase in transition distance relative to the sharp-leading-edge cylinder is in good agreement with the theoretical predictions of NACA Technical Report 1312. For relatively large values of the ratio of nose bluntness to transition distance, the effects of nose blunting were adverse for both the cone and the cylinder. In general, adverse effects due to blunting were larger for the flat bluntness than for the hemispherical or the round bluntness of equal bluntness height. Increasing the Mach number increased the size of bluntness required to induce adverse effects at constant free-stream Reynolds number per foot, delayed the adverse effects to higher values of Reynolds number per foot for constant nose bluntness, and reduced the abruptness of the transition decrease.

Monta, William J.↗

NASA NTRS1999

Cone-Probe Rake Design and Calibration for Supersonic Wind Tunnel Models

A series of experimental investigations were conducted at the NASA Langley Unitary Plan Wind Tunnel (UPWT) to calibrate cone-probe rakes designed to measure the flow field on 1-2% scale, high-speed wind tunnel models from Mach 2.15 to 2.4. The rakes were developed from a previous design that exhibited unfavorable measurement characteristics caused by a high probe spatial density and flow blockage from the rake body. Calibration parameters included Mach number, total pressure recovery, and flow angularity. Reference conditions were determined from a localized UPWT test section flow survey using a 10deg supersonic wedge probe. Test section Mach number and total pressure were determined using a novel iterative technique that accounted for boundary layer effects on the wedge surface. Cone-probe measurements were correlated to the surveyed flow conditions using analytical functions and recursive algorithms that resolved Mach number, pressure recovery, and flow angle to within +/-0.01, +/-1% and +/-0.1deg , respectively, for angles of attack and sideslip between +/-8deg. Uncertainty estimates indicated the overall cone-probe calibration accuracy was strongly influenced by the propagation of measurement error into the calculated results.

Won, Mark J.↗

NASA NTRS1990

Raked circular-cone aerobraking orbital transfer vehicle

An aerobraking orbital transfer vehicle (AOTV) (80) has aerobrake (82) with a blunted raked-off circular-cone configuration. The other components of the AOTV, including command/control module (95), fuel tanks (86, 88, 89 and 91), rocket engines (94) and afterbody (84), are positioned substantially along resultant force axis (104) of the AOTV (80). The axis (104) coincides with the resultant (sum of lift and drag) force vector. Afterbody (84) is mounted behind the aerobrake (82) with its length extending rearwardly from the aerobrake. The base flow clearance angle .phi. of the aerobrake (80) is 25.degree., thus allowing the afterbody (84) to extend rearwardly from the aerobrake (82) to a much greater extent than possible with a raked-off elliptic-cone aerobraking shield configuration. Afterbody size limitation and other problems associated with the raked-off elliptic-cone aerobraking shield configuration are alleviated by the combination of the aerobrake shape and positioning of the fuel tanks (86, 88, 89 and 91), rocket engines (94) and afterbody (84).

Park, Chul↗

NASA NTRS2008

Effects of Wall Cooling on Hypersonic Boundary Layer Receptivity Over a Cone

Effects of wall cooling on the receptivity process induced by the interaction of slow acoustic disturbances in the free-stream are numerically investigated for a boundary layer flow over a 5-degrees straight cone. The free-stream Mach number is 6.0 and the Reynolds number is 7.8x10(exp 6)/ft. Both the steady and unsteady solutions are obtained by solving the full Navier-Stokes equations using 5th-order accurate weighted essentially non-oscillatory (WENO) scheme for space discretization and using 3rd-order total variation diminishing (T VD) Runge-K utta scheme for time integration. Computations are performed for a cone with nose radius of 0.001 inch for adiabatic wall temperature (T(sub aw)), 0.75*T(sub aw), 0.5*T(sub aw), 0.40*T(sub aw), 0.30*T(sub aw), and 0.20*T(sub aw). Once the mean flow field is computed, disturbances are introduced at the upstream end of the computational domain. Generation of instability waves from leading edge region and receptivity of boundary layer to slow acoustic waves are investigated. Computations showed that wall cooling has strong stabilization effect on the first mode disturbances as was observed in the experiments. T ransition location moved to upstream when wall cooling was applied It is also found that the boundary layer is much more receptive to fast acoustic wave (by almost a factor of 50). When simulations performed using the same forcing frequency growth of the second mode disturbances are delayed with wall cooling and they attained values two times higher than that of adiabatic case. In 0.20*T(sub aw) case the transition Reynolds number is doubled compared to adiabatic conditions. The receptivity coefficient for adiabatic wall case (804 R) is 1.5225 and for highly cooled cones (241, and 161 R); they are in the order of 10(exp -3).

Kara, K.↗

NASA NTRS2008

SWCX Emission from the Helium Focusing Cone - Preliminary Results

Preliminary results from an XMM-Newton campaign to study solar wind charge exchange (SWCX) emission from the heliospheric focusing cone of interstellar helium are presented. The detections of enhanced O VII and O VIII emission from the cone are at the 2(sigma) and 4(sigma) levels. The solar wind charge exchange (SWCX) emission in the heliosphere not associated with distinct objects (e.g., comets and planets including exospheric material in and near Earth s magnetosheath) is proportional to the flux of the solar wind and the space density of neutral material. The neutral material originates in the interstellar medium (ISM) and passes through the solar system due to the relative motion of the Sun and the ISM. The flow of the neutral material through the solar system is strongly perturbed by the Sun both by gravity and by radiation pressure. Because of the relative radiative scattering cross sections and the effect of solar gravitation the density of interstellar hydrogen near the Sun is reduced while interstellar helium is gravitationally focused. This creates a helium focusing cone downstream of the Sun [e.g., 1, and references therein].

Snowden, S. L.↗

NASA NTRS2009

Aerothermodynamic Design of the Mars Science Laboratory Backshell and Parachute Cone

Aerothermodynamic design environments are presented for the Mars Science Laboratory entry capsule backshell and parachute cone. The design conditions are based on Navier-Stokes flowfield simulations on shallow (maximum total heat load) and steep (maximum heat flux) design entry trajectories from a 2009 launch. Transient interference effects from reaction control system thruster plumes were included in the design environments when necessary. The limiting backshell design heating conditions of 6.3 W/sq cm for heat flux and 377 J/sq cm for total heat load are not influenced by thruster firings. Similarly, the thrusters do not affect the parachute cover lid design environments (13 W/sq cm and 499 J/sq cm). If thruster jet firings occur near peak dynamic pressure, they will augment the design environments at the interface between the backshell and parachute cone (7 W/sq cm and 174 J/sq cm). Localized heat fluxes are higher near the thruster fairing during jet firings, but these areas did not require additional thermal protection material. Finally, heating bump factors were developed for antenna radomes on the parachute cone

Edquist, Karl T.↗

NASA NTRS2010

Camera Layout Design for the Upper Stage Thrust Cone

Engineers in the Integrated Design and Analysis Division (EV30) use a variety of different tools to aid in the design and analysis of the Ares I vehicle. One primary tool in use is Pro-Engineer. Pro-Engineer is a computer-aided design (CAD) software that allows designers to create computer generated structural models of vehicle structures. For the Upper State thrust cone, Pro-Engineer was used to assist in the design of a layout for two camera housings. These cameras observe the separation between the first and second stage of the Ares I vehicle. For the Ares I-X, one standard speed camera was used. The Ares I design calls for two separate housings, three cameras, and a lighting system. With previous design concepts and verification strategies in mind, a new layout for the two camera design concept was developed with members of the EV32 team. With the new design, Pro-Engineer was used to draw the layout to observe how the two camera housings fit with the thrust cone assembly. Future analysis of the camera housing design will verify the stability and clearance of the camera with other hardware present on the thrust cone.

Wooten, Tevin↗

NASA NTRS2011

Experimental Study of Second-Mode Instabilities on a 7-Degree Cone at Mach 6

Experiments have been carried out in the NASA Langley Research Center 20-Inch Mach 6 Air Tunnel to measure the second-mode boundary-layer instability on a 7deg half-angle cone using high-frequency pressure sensors. Data were obtained with both blunt and sharp nosetips installed on the cone. The second-mode wave amplitudes were observed to saturate and then begin to decrease in the Langley tunnels, indicating wave breakdown. Pressure fluctuation measurements and thermocouple data indicated the location of transition along the cone at the different conditions tested. Comparisons between the power density spectra obtained during the current test and previous data from the Langley 15-Inch Mach 6 High Temperature Tunnel and the Boeing/AFOSR Mach 6 Quiet tunnel illustrate the effect of tunnel noise on instability growth and transition.

Rufer, Shann J.↗

NASA NTRS2016

Frequency-Range Distribution of Boulders Around Cone Crater: Relevance to Landing Site Hazard Avoidance

Boulders represent a landing hazard that must be addressed in the planning of future landings on the Moon. A boulder under a landing leg can contribute to deck tilt and boulders can damage spacecraft during landing. Using orbital data to characterize boulder populations at locations where landers have safely touched down (Apollo, Luna, Surveyor, and Chang'e-3 sites) is important for determining landing hazard criteria for future missions. Additionally, assessing the distribution of boulders can address broader science issues, e.g., how far craters distribute boulders and how this distribution varies as a function of crater size and age. The availability of new Lunar Reconnaissance Orbiter Camera (LROC) Narrow Angle Camera (NAC) images [1] enables the use of boulder size- and range frequency distributions for a variety of purposes [2-6]. Boulders degrade over time and primarily occur around young or fresh craters that are large enough to excavate bedrock. Here we use NAC images to analyze boulder distributions around Cone crater (340 m diameter) at the Apollo 14 site. Cone crater (CC) was selected because it is the largest crater where astronaut surface photography is available for a radial traverse to the rim. Cone crater is young (approximately 29 Ma [7]) relative to the time required to break down boulders [3,8], giving us a data point for boulder range-frequency distributions (BRFDs) as a function of crater age.

Clegg-Watkins, R. N.↗