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At least 289 records · Page 16
All sourcesNASA NTRSDOE OSTI
NASA NTRS2006

Understanding Columbia's Reentry Problems

Soon after the Space Shuttle Columbia accident occurred last year, a group of CFD analysts from NASA centers and private industry was organized to help determine the cause of the accident. This group was under the direction of the Applied Aeroscience and CFD Branch of the Aeroscience and Flight Mechanics Division at the Johnson Space Center. For external flow simulations, noncommercia2 CFD codes that specialize in hypersonic or high Mach number flows were used. These tools were used to determine heating rates, pressures, and temperatures for a large number of vehicle damage scenarios. Lockheed Martin Space Operations was called upon to provide CFD support in the area of internal flows within the shuttle wing cavity, and for these simulations, FLUENT 6.1 was chosen. Two large-scale, simplified models were m to understand the flow patterns once a breach of the internal wing cavity was initiated. The results were primarily used to visualize flow patterns within the wing cavity. The first CFD model included the entire lee wing without the wheel well cavity. The purpose of the first model, which did not include the reinforced carbon-carbon (RCC) cavity along the wing leading edge, was to visualize the flow field within the wing cavity immediately after the leading edge spar breach, This model assumed that the flow coming into the wing cavity was normal to the spar. It included all of the primary vents that allow for flow between the main cavities of the wing. A six-inch diameter hole was modeled in the spar at the approximate location where the spar breach was judged to have occurred, which was between RCC panels 8 and 9. The results of the modeling showed that at this location, the high temperature, high velocity gas stream entering the wing cavity impinged on the outboard wheel well cavity. Instrumentation in the Shuttle wheel well cavity registered abnormal temperatures during reentry, so the FLUENT results helped support the conclusion of the accident investigation team that the spar breach was in the RCC panel 8-9 area, and that the initial spar breach was likely entering the wing cavity normal to the spar. This model also showed that the flow entering the wing cavity tended to swirl within the cavity just outboard of the wheel well, and did not initially penetrate further into the rear cavities of the wing. The second CFD model was a 2-D simulation of the left wing cavity and the RCC cavity. It was used to visualize the flow through the RCC breach, through the wing spar breach, and into the wing cavity directly outboard of the wheel well. The purpose of this model was to verify whether or not it was possible for the flow to come into the wing cavity normal to the leading edge spar. The results from this 2-D model showed that the internal structure behind the RCC panels probably deflected the flow entering the RCC cavity so that it impinged normal to the spar. As in the 3-D model, this deflected flow stream was Found to impinge on the wheel well outer wall. This model again supported the conclusions regarding the location of the spar breach and how the flow behaved inside the wing cavity.

Paul, Thomas↗
NASA NTRS2004

Understanding Columbia's Reentry Problems

Soon after the Space Shuttle Columbia accident occurred last year, a group of CFD analysts from NASA centers and private industry was organized to help determine the cause of the accident. This group was under the direction of the Applied Aeroscience and CFD Branch of the Aeroscience and Flight Mechanics Division at the Johnson Space Center. For external flow simulations, noncommercial CFD codes that specialize in hypersonic or high Mach number flows were used. These tools were used to determine heating rates, pressures, and temperatures for a large number of vehicle damage scenarios. Lockheed Martin Space Operations was called upon to provide CFD support in the area of internal flows within the shuttle wing cavity, and or these simulations, FLUENT 6.1 was chosen. Two large-scale, simplified models were run to understand the flow patterns once a breach of the internal wing cavity was initiated. The results were primarily used to visualize flow patterns within the wing cavity. The first CFD model included the entire left wing without the wheel. well cavity. The purpose of the first model, which did not include the reinforced carbon-carbon (RCC) cavity along the wing leading edge, was to visualize the flow field within the wing cavity immediately after the leading edge spar breach, This model assumed that the flow coming into the wing cavity was nom1 to the spar. It included all o f the primary vents that allow for flow between the main cavities of the wing. A six-inch diameter hole was modeled in the spar at the approximate location where the spar breach was judged to have occurred, which was between RCC panels 8 and 9. The results of the modeling showed that at this location, the high temperature, high velocity gas stream entering the wing cavity impinged on the outboard wheel well cavity. Instrumentation in the Shuttle wheel well cavity registered abnormal temperatures during reentry, so the FLUENT results helped support the conclusion of the accident investigation team that the spar breach was in the RCC panel 8-9 area, and that the initial spar breach was likely entering the wing cavity normal to the spar. This model also showed that the flow entering the wing cavity tended to swirl within the cavity just outboard of the wheel well, and did not initially penetrate further into the rear cavities of the wing. The second CFD model was a 2-D simulation of the left wing cavity and the RCC cavity. It was used to visualize the flow through the RCC breach, through the wing spar breach, and into the wing cavity directly outboard of the wheel well. The purpose of this model was to verify whether or not it was possible for the flow to come into the wing cavity normal to the leading edge spar. The results from this 2-D model showed that the internal structure behind the RCC panels probably deflected the flow entering the RCC cavity so that it impinged normal to the spar. As in the 3-D model, this deflected flow stream was found to impinge on the wheel well outer wall. This model again supported the conclusions regarding the location of the spar breach and how the flow behaved inside the wing cavity.

Paul, Thomas H.↗
NASA NTRS1992

A unified approach for numerical simulation of viscous compressible and incompressible flows over adiabatic and isothermal walls

A new formulation (including the choice of variables, their non-dimensionalization, and the form of the artificial viscosity) is proposed for the numerical solution of the full Navier-Stokes equations for compressible and incompressible flows with heat transfer. With the present approach, the same code can be used for constant as well as variable density flows. The changes of the density due to pressure and temperature variations are identified and it is shown that the low Mach number approximation is a special case. At zero Mach number, the density changes due to the temperature variation are accounted for, mainly through a body force term in the momentum equation. It is also shown that the Boussinesq approximation of the buoyancy effects in an incompressible flow is a special case. To demonstrate the new capability, three examples are tested. Flows in driven cavities with adiabatic and isothermal walls are simulated with the same code as well as incompressible and supersonic flows over a wall with and without a groove. Finally, viscous flow simulations of an oblique shock reflection from a flat plate are shown to be in good agreement with the solutions available in literature.

Hafez, M.↗
NASA NTRS2011

Effects of Cavities and Protuberances on Transition over Hypersonic Vehicles

Surface protuberances and cavities on a hypersonic vehicle are known to cause several aerodynamic or aerothermodynamic issues. Most important of all, premature transition due to these surface irregularities can lead to a significant rise in surface heating. To help understand laminar-turbulent transition induced by protuberances or cavities on a Crew Exploration Vehicle (CEV) surface, high-fidelity numerical simulations are carried out for both types of trips on a CEV wind tunnel model. Due to the large bluntness, these surface irregularities reside in an accelerating subsonic boundary layer. For the Mach 6 wind tunnel conditions with a roughness Reynolds number Re(sub kk) of 800, it was found that a protuberance with a height to boundary layer thickness ratio of 0.73 leads to strong wake instability and spontaneous vortex shedding, while a cavity with identical geometry only causes a rather weak flow unsteadiness. The same cavity with a larger Reynolds number also leads to similar spontaneous vortex shedding and wake instability. The wake development and the formation of hairpin vortices for both protuberance and cavity were found to be qualitatively similar to that observed for an isolated hemisphere submerged in a subsonic, low speed flat-plate boundary layer. However, the shed vortices and their accompanying instability waves were found to be slightly stabilized downstream by the accelerating boundary layer along the CEV surface. Despite this stabilizing influence, it was found that the wake instability spreads substantially in both wall-normal and azimuthal directions as the flow is evolving towards a transitional state. Similarities and differences between the wake instability behind a protuberance and a cavity are investigated. Computations for the Mach 6 boundary layer over a slender cylindrical roughness element with a height to the boundary layer thickness of about 1.1 also shows spontaneous vortex shedding and strong wake instability. Comparisons of detailed flow structures associated with protuberances at subsonic and supersonic edge Mach numbers indicate distinctively different instability mechanisms.

Chang, Chau-Lyan↗
NASA NTRS2002

Lattice Boltzmann Method for 3-D Flows with Curved Boundary

In this work, we investigate two issues that are important to computational efficiency and reliability in fluid dynamics applications of the lattice, Boltzmann equation (LBE): (1) Computational stability and accuracy of different lattice Boltzmann models and (2) the treatment of the boundary conditions on curved solid boundaries and their 3-D implementations. Three athermal 3-D LBE models (D3QI5, D3Ql9, and D3Q27) are studied and compared in terms of efficiency, accuracy, and robustness. The boundary treatment recently developed by Filippova and Hanel and Met et al. in 2-D is extended to and implemented for 3-D. The convergence, stability, and computational efficiency of the 3-D LBE models with the boundary treatment for curved boundaries were tested in simulations of four 3-D flows: (1) Fully developed flows in a square duct, (2) flow in a 3-D lid-driven cavity, (3) fully developed flows in a circular pipe, and (4) a uniform flow over a sphere. We found that while the fifteen-velocity 3-D (D3Ql5) model is more prone to numerical instability and the D3Q27 is more computationally intensive, the 63Q19 model provides a balance between computational reliability and efficiency. Through numerical simulations, we demonstrated that the boundary treatment for 3-D arbitrary curved geometry has second-order accuracy and possesses satisfactory stability characteristics.

Mei, Renwei↗
NASA NTRS1996

Effect of Cavity Injection on Wall Drag in a Supersonic Flow with Reflecting Shocks

The wall drag test tunnel at NASA Langley Research Center was used to evaluate simulated scramjet fuel injection into a wall cavity. In this tunnel, one wall consists of interchangeable aluminum plates attached to an air bearing suspension system. The plates were equipped with load cells to measure drag forces and static taps to determine pressure distributions. The plates were exposed to a Mach 2 air stream at a total pressure of 115 psia (793 kPa). This flow field contained a train of weak unsteady, reflecting shock waves that were produced in the nozzle assembly located upstream of the test section.

Castiglone, L. A.↗
NASA NTRSDate not supplied

Shuttle Main Propulsion System LH2 Feed Line and Inducer Simulations

This viewgraph presentation includes simulations of the unsteady flow field in the LH2 feed line, flow line, flow liner, backing cavity and inducer of Shuttle engine #1. It also evaluates aerodynamic forcing functions which may contribute to the formation of the cracks observed on the flow liner slots. The presentation lists the numerical methods used, and profiles a benchmark test case.

Daniel Dorney↗
NASA NTRS2005

Analysis of Coupled Seals, Secondary and Powerstream Flow Fields in Aircraft and Aerospace Turbomachines

Higher power, high efficiency gas turbine engines require optimization of the seals and secondary flow systems as well as their impact on the powerstream. This work focuses on two aspects: 1. To apply the present day CFD tools (SCISEAL) to different real-life secondary flow applications from different original equipment manufacturers (OEM s) to provide feedback data and 2. Develop a computational methodology for coupled time-accurate simulation of the powerstream and secondary flow with emphasis on the interaction between the disk-cavity and rim seals flows with the powerstream (SCISEAL-MS-TURBO). One OEM simulation was of the Allison Engine Company T-56 turbine drum cavities including conjugate heat transfer with good agreement with data and provided design feedback information. Another was the GE aspirating seal where the 3-D CFD simulations played a major role in analysis and modification of that seal configuration. The second major objective, development of a coupled flow simulation capability was achieved by using two codes MS-TURBO for the powerstream and SCISEAL for the secondary flows with an interface coupling algorithm. The coupled code was tested against data from three differed configurations: 1. bladeless-rotor-stator-cavity turbine test rig, 2. UTRC high pressure turbine test rig, and, 3. the NASA Low-Speed-Air Compressor rig (LSAC) with results and limitations discussed herein.

Athavale, M. M.↗
NASA NTRS1985

Supercritical airfoil drag reduction by passive shock wave/boundary layer control in the Mach number range .75 to .90

Airfoils operating in the transonic region are subject to large increases in drag due to shock wave/boundary layer interactions. The concept of passive shock wave/boundary layer control seeks drag reduction by placing a thin cavity with a porous top surface at the airfoil chordwise position where a shock wave would normally occur. The higher pressure behind the shock wave circulates flow through the cavity to the lower pressure ahead of the shock wave. The effects from this circulation prevent boundary layer separation and reduce entropy increases through the shock wave. In this investigation this concept is studied at a freestream Mach number range of .75 and .90. The Mach number distributions over the model, the wake impact pressure surveys used to determine profile drag and schlieren photographs for 2.8 percent porosity and solid airfoil cases are presented and compared. Results indicate that the profile drag coefficient can be reduced by as much as 40 percent through the use of this passive drag control system.

Nagamatsu, H. T.↗
NASA NTRS1978

Petrogenesis of melt rocks, Manicouagan impact structure, Quebec

It is suggested, on the basis of previous theoretical studies of shock waves, that the Manicouagan melt formed in 1 or 2 s in a 5-km-radius hemisphere near the point of impact. The melt and the less shocked debris surrounding it flowed downward and outward for a few minutes until the melt formed a lining of a 5- to 8-km deep, 15- to 22-km-radius cavity. Extremely turbulent flow thoroughly homogenized the melt and promoted the incorporation and progressive digestion of debris that had been finely fragmented (but not melted) to grain sizes of less than one mm by the passage of the shock waves. The equilibration of clasts and melt, plagioclase nucleation, and readjustment of the crater floor are discussed.

Simonds, C. H.↗
NASA NTRS1995

Evaluation of Aircraft Platforms for SOFIA by Computational Fluid Dynamics

The selection of an airborne platform for the Stratospheric Observatory for Infrared Astronomy (SOFIA) is based not only on economic cost, but technical criteria, as well. Technical issues include aircraft fatigue, resonant characteristics of the cavity-port shear layer, aircraft stability, the drag penalty of the open telescope bay, and telescope performance. Recently, two versions of the Boeing 747 aircraft, viz., the -SP and -200 configurations, were evaluated by computational fluid dynamics (CFD) for their suitability as SOFIA platforms. In each configuration the telescope was mounted behind the wings in an open bay with nearly circular aperture. The geometry of the cavity, cavity aperture, and telescope was identical in both platforms. The aperture was located on the port side of the aircraft and the elevation angle of the telescope, measured with respect to the vertical axis, was 500. The unsteady, viscous, three-dimensional, aerodynamic and acoustic flow fields in the vicinity of SOFIA were simulated by an implicit, finite-difference Navier-Stokes flow solver (OVERFLOW) on a Chimera, overset grid system. The computational domain was discretized by structured grids. Computations were performed at wind-tunnel and flight Reynolds numbers corresponding to one free-stream flow condition (M = 0.85, angle of attack alpha = 2.50, and sideslip angle beta = 0 degrees). The computational domains consisted of twenty-nine(29) overset grids in the wind-tunnel simulations and forty-five(45) grids in the simulations run at cruise flight conditions. The maximum number of grid points in the simulations was approximately 4 x 10(exp 6). Issues considered in the evaluation study included analysis of the unsteady flow field in the cavity, the influence of the cavity on the flow across empennage surfaces, the drag penalty caused by the open telescope bay, and the noise radiating from cavity surfaces and the cavity-port shear layer. Wind-tunnel data were also available to compare to the CFD results; the data permitted an assessment of CFD as a design tool for the SOFIA program.

Klotz, S. P.↗
NASA NTRS1988

Numerical analysis of laminar and turbulent incompressible flows using the finite element Fluid Dynamics Analysis Package (FIDAP)

The purpose of the study is the evaluation of the numerical accuracy of FIDAP (Fluid Dynamics Analysis Package). Accordingly, four test problems in laminar and turbulent incompressible flows are selected and the computational results of these problems compared with other numerical solutions and/or experimental data. These problems include: (1) 2-D laminar flow inside a wall-driven cavity; (2) 2-D laminar flow over a backward-facing step; (3) 2-D turbulent flow over a backward-facing step; and (4) 2-D turbulent flow through a turn-around duct.

Sohn, Jeong L.↗
NASA NTRS1995

Flowfield and acoustic characteristics of telescope cavity in SOFIA platform

Unsteady three-dimensional flowfields are calculated for the Stratospheric Observatory For Infrared Astronomy (SOFIA) at both free-flight cruise and wind tunnel conditions with a view to help in the design process of an acoustically quiet telescope cavity and to understand the flow physics of a three dimensional cavity. The calculation method is based on the numerical solution of thin layer Navier-Stokes equations on a Chimera overset grid system. The Boeing 747-200 aircraft is examined as one option for the SOFIA platform. The flowfield domain is composed of 45 grids consisting of over 4.1 million points. Numerical simulations are performed for both wind tunnel and free-flight cruise conditions at one freestream condition of M(infinity) = 0.85, alpha = 2.5 deg. Comparison of results from wind tunnel simulation show good agreement with experimental data for time-averaged surface pressures, drag for the empennage, and sound pressure levels and power spectra at various locations within the cavity and on the telescope. The presence of the open cavity induces an incremental drag increase, an increased acoustic radiation, and an increase in unsteady pressure loads on the telescope. Its impact on the effectiveness of aircraft control surfaces appears minimal.

Srinivasan, G. R.↗
NASA NTRS2009

Characterization of a Plasmoid in the Afterglow of a Supersonic Flowing Microwave Discharge

We performed a detailed characterization a plasmoid in the afterglow region of an Ar supersonic microwave cavity discharge. The supersonic flow was generated using a convergent-divergent nozzle upstream of the discharge region. A cylindrical cavity was used to sustain a discharge in the pressure range of 100-600 Pa. Optical emission spectroscopy was used to observe populations of excited and ionic species in the plasmoid region. Plasmoid formation in the supersonic flowing afterglow located downstream from the primary microwave cavity discharge was characterized by measuring the radial and axial distributions of Argon excited states and Argon ions. More experiments are being carried out on the plasmoid to understand the discharge parameters within the region, i.e. rotational temperature, vibrational temperature, electron density, and how the electrodynamic and aerodynamic effects combine to form this plasmoid.

Drake, D. J.↗
NASA NTRS1982

Store separation from cavities at supersonic flight speeds

An experimental investigation has been conducted to determine the aerodynamic characteristics of a typical wing-control missile configuration during separation from a box-type cavity having depth to length ratios (D/L) ranging from 0.088 to 0.225. The cavity was located in a splitter plate that spanned the low Mach number test section of the Langley Unitary Plan Wind Tunnel. Aerodynamic characteristics are presented for Mach 2.36 and a free-stream unit Reynolds number of 2,000,000/ft. For the shallow cavity (D/L = 0.088), large interactions existed between the cavity and the flat plate flow field, which resulted in unfavorable separation characteristics for the missile model. For the deep cavity (D/L = 0.225), the flat plate flow field essentially bridged the cavity, resulting in minor interactions and favorable separation characteristics for the missile model.

Stallings, R. L., Jr.↗
NASA NTRS1975

Visual study of the effect of grazing flow on the oscillatory flow in a resonator orifice

Grazing flow and oscillatory flow in an orifice were studied in a plexiglass flow channel with a single side branch Helmholtz resonator using water as the fluid medium. An oscillatory flow was applied to the resonatory cavity, and color dyes were injected in both the orifice and the grazing flow field to record the motion of the fluid. The flow regimes associated with linear and nonlinear (high sound pressure level) impedances with and without grazing flows were recorded by a high-speed motion-picture camera. Appreciable differences in the oscillatory flow field were seen in the various flow regimes. With high grazing flows, the outflow and inflow from the resonator cavity are found to be asymmetric. The visual study confirms that jet energy loss during flow into a resonator cavity is much larger than the loss for ejection from the cavity into the grazing flow. For inflow into the resonator cavity, the effective orifice area was significantly reduced.

Baumeister, K. J.↗
NASA NTRS1987

Independent Orbiter Assessment (IOA): Analysis of the purge, vent and drain subsystem

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. This report documents the independent analysis results corresponding to the Orbiter PV and D (Purge, Vent and Drain) Subsystem hardware. The PV and D Subsystem controls the environment of unpressurized compartments and window cavities, senses hazardous gases, and purges Orbiter/ET Disconnect. The subsystem is divided into six systems: Purge System (controls the environment of unpressurized structural compartments); Vent System (controls the pressure of unpressurized compartments); Drain System (removes water from unpressurized compartments); Hazardous Gas Detection System (HGDS) (monitors hazardous gas concentrations); Window Cavity Conditioning System (WCCS) (maintains clear windows and provides pressure control of the window cavities); and External Tank/Orbiter Disconnect Purge System (prevents cryo-pumping/icing of disconnect hardware). Each level of hardware was evaluated and analyzed for possible failure modes and effects. Criticality was assigned based upon the severity of the effect for each failure mode. Four of the sixty-two failure modes analyzed were determined as single failures which could result in the loss of crew or vehicle. A possible loss of mission could result if any of twelve single failures occurred. Two of the criticality 1/1 failures are in the Window Cavity Conditioning System (WCCS) outer window cavity, where leakage and/or restricted flow will cause failure to depressurize/repressurize the window cavity. Two criticality 1/1 failures represent leakage and/or restricted flow in the Orbiter/ET disconnect purge network which prevent cryopumping/icing of disconnect hardware. Each level of hardware was evaluated and analyzed for possible failure modes and effects. Criticality was assigned based upon the severity of the effect for each failure mode.

Bynum, M. C., III↗
NASA NTRS2000

Direct Numerical Simulation of Automobile Cavity Tones

The Navier Stokes equation is solved computationally by the Dispersion-Relation-Preserving (DRP) scheme for the flow and acoustic fields associated with a laminar boundary layer flow over an automobile door cavity. In this work, the flow Reynolds number is restricted to R(sub delta*) < 3400; the range of Reynolds number for which laminar flow may be maintained. This investigation focuses on two aspects of the problem, namely, the effect of boundary layer thickness on the cavity tone frequency and intensity and the effect of the size of the computation domain on the accuracy of the numerical simulation. It is found that the tone frequency decreases with an increase in boundary layer thickness. When the boundary layer is thicker than a certain critical value, depending on the flow speed, no tone is emitted by the cavity. Computationally, solutions of aeroacoustics problems are known to be sensitive to the size of the computation domain. Numerical experiments indicate that the use of a small domain could result in normal mode type acoustic oscillations in the entire computation domain leading to an increase in tone frequency and intensity. When the computation domain is expanded so that the boundaries are at least one wavelength away from the noise source, the computed tone frequency and intensity are found to be computation domain size independent.

Kurbatskii, Konstantin↗