SEARCH · Search NASA
Results for “Plume flow”
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.
Single engine: Plume induced flow separation
Explore the source record for details and available documents.
Numerical solution of axisymmetric boattail flow fields with plume simulators
Turbulent separating flows over axisymmetric afterbody-boattail configurations with solid sting plume simulators are computed with a time-dependent finite-difference method to solve the compressible Navier-Stokes equations. The Reynolds stress terms are replaced with a two-layer eddy viscosity model including a relaxation formula to model the nonequilibrium effects of the separated flow. The mesh is alined with the boattail body through an analytic transformation which accommodates a wide variety of boattail geometries. Numerical results for a series of boattail geometries over a wide range of Reynolds number (140,000 to 140 million) are presented and, when possible, compared with experimental data or independent numerical results.
Development of Computational Aeroacoustics Code for Jet Noise and Flow Prediction
Accurate prediction of jet fan and exhaust plume flow and noise generation and propagation is very important in developing advanced aircraft engines that will pass current and future noise regulations. In jet fan flows as well as exhaust plumes, two major sources of noise are present: large-scale, coherent instabilities and small-scale turbulent eddies. In previous work for the NASA Glenn Research Center, three strategies have been explored in an effort to computationally predict the noise radiation from supersonic jet exhaust plumes. In order from the least expensive computationally to the most expensive computationally, these are: 1) Linearized Euler equations (LEE). 2) Very Large Eddy Simulations (VLES). 3) Large Eddy Simulations (LES). The first method solves the linearized Euler equations (LEE). These equations are obtained by linearizing about a given mean flow and the neglecting viscous effects. In this way, the noise from large-scale instabilities can be found for a given mean flow. The linearized Euler equations are computationally inexpensive, and have produced good noise results for supersonic jets where the large-scale instability noise dominates, as well as for the tone noise from a jet engine blade row. However, these linear equations do not predict the absolute magnitude of the noise; instead, only the relative magnitude is predicted. Also, the predicted disturbances do not modify the mean flow, removing a physical mechanism by which the amplitude of the disturbance may be controlled. Recent research for isolated airfoils' indicates that this may not affect the solution greatly at low frequencies. The second method addresses some of the concerns raised by the LEE method. In this approach, called Very Large Eddy Simulation (VLES), the unsteady Reynolds averaged Navier-Stokes equations are solved directly using a high-accuracy computational aeroacoustics numerical scheme. With the addition of a two-equation turbulence model and the use of a relatively coarse grid, the numerical solution is effectively filtered into a directly calculated mean flow with the small-scale turbulence being modeled, and an unsteady large-scale component that is also being directly calculated. In this way, the unsteady disturbances are calculated in a nonlinear way, with a direct effect on the mean flow. This method is not as fast as the LEE approach, but does have many advantages to recommend it; however, like the LEE approach, only the effect of the largest unsteady structures will be captured. An initial calculation was performed on a supersonic jet exhaust plume, with promising results, but the calculation was hampered by the explicit time marching scheme that was employed. This explicit scheme required a very small time step to resolve the nozzle boundary layer, which caused a long run time. Current work is focused on testing a lower-order implicit time marching method to combat this problem.
Molecular gas dynamics applied to low-thrust propulsion
The Direct Simulation Monte Carlo method is currently being applied to study flowfields of small thrusters, including both the internal nozzle and the external plume flow. The DSMC method is employed because of its inherent ability to capture nonequilibrium effects and proper boundary physics in low-density flow that are not readily obtained by continuum methods. Accurate prediction of both the internal and external nozzle flow is important in determining plume expansion which, in turn, bears directly on impingement and contamination effects.
High Energy Boundary Conditions for a Cartesian Mesh Euler Solver
Inlets and exhaust nozzles are often omitted or fared over in aerodynamic simulations of aircraft due to the complexities involving in the modeling of engine details such as complex geometry and flow physics. However, the assumption is often improper as inlet or plume flows have a substantial effect on vehicle aerodynamics. A tool for specifying inlet and exhaust plume conditions through the use of high-energy boundary conditions in an established inviscid flow solver is presented. The effects of the plume on the flow fields near the inlet and plume are discussed.
Laboratory behavior of a plasma plume injected into the magnetized plasma flow
Plasma plumes are injected into the tail of the simulated earth's magnetosphere produced by an interaction between the simulated solar wind and a magnetic dipole. The behavior of laboratory artificial plasma plumes injected into the magnetized plasma flow is discussed in conjunction with the AMPTE artificial comet experiments (Minami et al., 1986) and other active chemical release experiments in space.
Space shuttle vehicle rocket plume impingement study for separation analysis. Tasks 2 and 3: Definition and preliminary plume impingement analysis for the MSC booster
The results are presented of a space shuttle plume impingement study for the Manned Spacecraft Center configuration. This study was conducted as two tasks which were to (1) define the orbiter main stage engine exhaust plume flow field, and (2) define the plume impingement heating, force and resulting moment environments on the booster during the staging maneuver. To adequately define these environments during the staging maneuver and allow for deviation from the nominal separation trajectory, a multitude of relative orbiter/booster positions are analyzed which map the region that contains the separation trajectories. The data presented can be used to determine a separation trajectory which will result in acceptable impingement heating rates, forces, and the resulting moments. The data, presented in graphical form, include the effect of roll, pitch and yaw maneuvers for the booster. Quasi-steady state analysis methods were used with the orbiter engine operating at full thrust. To obtain partial thrust results, simple ratio equations are presented.
Experimental results for a hypersonic nozzle/afterbody flow field
The flow field created by the interaction of a single-expansion-ramp-nozzle (SERN) flow with a hypersonic external stream has been experimentally characterized using a generic nozzle/afterbody model in the 3.5-foot hypersonic wind tunnel of the NASA Ames Research Center. The presented results include oil-flow and shadowgraph flow visualization photographs, afterbody surface-pressure distributions, boundary layer rake measurements, and Preston-tube skin-friction measurements. The design, construction, and operation of the model was found to be successful. Surface oil-flow patterns show that the jet-plume flow attaches to the afterbody surface at jet pressure ratios between 154 and 234. The oil flow also shows the pattern of lines where the jet flow separates from the ramp, apparently as a result of interaction of the jet-plume internal shock wave with the ramp boundary layer.
Validation of Inlet and Exhaust Boundary Conditions for a Cartesian Method
Inlets and exhaust nozzles are often omitted in aerodynamic simulations of aircraft due to the complexities involved in the modeling of engine details and flow physics. However, the omission is often improper since inlet or plume flows may have a substantial effect on vehicle aerodynamics. A method for modeling the effect of inlets and exhaust plumes using boundary conditions within an inviscid Cartesian flow solver is presented. This approach couples with both CAD systems and legacy geometry to provide an automated tool suitable for parameter studies. The method is validated using two and three-dimensional test problems which are compared with both theoretical and experimental results. The numerical results demonstrate excellent agreement with theory and available data, even for extremely strong jets and very sensitive inlets.
Space shuttle contamination due to backflow from control motor exhaust
Spacecraft contamination of the space shuttle orbiter and accompanying Spacelab payloads is studied. The scattering of molecules from the vernier engines and flash evaporator nozzle after impingement on the orbiter wing surfaces, and the backflow of molecules out of the flash evaporator nozzle plume flow field due to intermolecular collisions in the plume are the problems discussed. A method was formulated for dealing with these problems, and detailed results are given.
Optical Diagnostic Imaging of Multi-Rocket Plume-Induced Base Flow Environments
Non-intrusive optical diagnostic imaging was used for the first time to visualize multi-rocket plume-induced reacting base flows to simulate launch vehicle ascent from sea-level to 250,000 ft. In particular, planar laser induced florescence (PLIF) and infrared (IR) imaging were implemented for the first time to visualize and quantify base flow and rocket plume environments from sub-scale, short-duration propulsion models within a shock tunnel facility. This report discusses the successful imaging diagnostic methods for capturing base flow features and dynamics as a function of altitude. Important base flow and plume features were captured with PLIF and IR diagnostics to develop a conceptual base flow physics model. This imaging data specifically provides insight into the Space Launch System vehicle core-stage and Exploration Upper Stage base environments and further validates short-duration ground test techniques and computational modeling.
Orbital Maneuvering Vehicle (OMV) plume and plume effects study
The objective was to characterize the Orbital Maneuvering Vehicle (OMV) propulsion and attitude control system engine exhaust plumes and predict the resultant plume impingement pressure, heat loads, forces, and moments. Detailed description is provided of the OMV gaseous nitrogen (GN2) thruster exhaust plume flow field characteristics calculated with the RAMP2 snd SFPGEN computer codes. Brief descriptions are included of the two models, GN2 thruster characteristics and RAMP2 input data files. The RAMP2 flow field could be recalculated by other organizations using the information presented. The GN2 flow field can be readily used by other organizations who are interested in GN2 plume induced environments which require local flow field properties which can be supplied using the SFPGEN GN2 model.
Rocket plume properties measured in space simulators
Molecular sink facility and 25-foot space simulator have been used to distinguish nature of exhaust plumes from nozzles with relatively large internal boundary layer flow. Plume density has been measured by electron beam/photomultiplier system.
Assessment of analytical and experimental techniques utilized in conducting plume technology tests 575 and 593
Since exhaust plumes affect vehicle base environment (pressure and heat loads) and the orbiter vehicle aerodynamic control surface effectiveness, an intensive program involving detailed analytical and experimental investigations of the exhaust plume/vehicle interaction was undertaken as a pertinent part of the overall space shuttle development program. The program, called the Plume Technology program, has as its objective the determination of the criteria for simulating rocket engine (in particular, space shuttle propulsion system) plume-induced aerodynamic effects in a wind tunnel environment. The comprehensive experimental program was conducted using test facilities at NASA's Marshall Space Flight Center and Ames Research Center. A post-test examination of some of the experimental results obtained from NASA-MSFC's 14 x 14-inch trisonic wind tunnel is presented. A description is given of the test facility, simulant gas supply system, nozzle hardware, test procedure and test matrix. Analysis of exhaust plume flow fields and comparison of analytical and experimental exhaust plume data are presented.
Determining Aerosol Plume Height from Two GEO Imagers: Lessons from MISR and GOES
Aerosol plume height is a key parameter to determine impacts of particulate matters generated from biomass burning, wind-blowing dust, and volcano eruption. Retrieving cloud top height from stereo imageries from two GOES (Geostationary Operational Environmental Satellites) have been demonstrated since 1970's and the principle should work for aerosol plumes if they are optically thick. The stereo technique has also been used by MISR (Multiangle Imaging SpectroRadiometer) since 2000 that has nine look angles along track to provide aerosol height measurements. Knowing the height of volcano aerosol layers is as important as tracking the ash plume flow for aviation safety. Lack of knowledge about ash plume height during the 2010 Eyja'rjallajokull eruption resulted in the largest air-traffic shutdown in Europe since World War II. We will discuss potential applications of Asian GEO satellites to make stereo measurements for dust and volcano plumes.
Supersonic Flow of Chemically Reacting Gas-Particle Mixtures. Volume 2: RAMP - A Computer Code for Analysis of Chemically Reacting Gas-Particle Flows
A computer program written in conjunction with the numerical solution of the flow of chemically reacting gas-particle mixtures was documented. The solution to the set of governing equations was obtained by utilizing the method of characteristics. The equations cast in characteristic form were shown to be formally the same for ideal, frozen, chemical equilibrium and chemical non-equilibrium reacting gas mixtures. The characteristic directions for the gas-particle system are found to be the conventional gas Mach lines, the gas streamlines and the particle streamlines. The basic mesh construction for the flow solution is along streamlines and normals to the streamlines for axisymmetric or two-dimensional flow. The analysis gives detailed information of the supersonic flow and provides for a continuous solution of the nozzle and exhaust plume flow fields. Boundary conditions for the flow solution are either the nozzle wall or the exhaust plume boundary.
Application of Background Oriented Schlieren for Altitude Testing of Rocket Engines
A series of experiments was performed to determine the feasibility of using the Background Oriented Schlieren, BOS, flow visualization technique to image a simulated, small, rocket engine, plume under altitude test conditions. Testing was performed at the NASA Glenn Research Centers Altitude Combustion Stand, ACS, using nitrogen as the exhaust gas simulant. Due to limited optical access to the facility test capsule, all of the hardware required to conduct the BOS were located inside the vacuum chamber. During the test series 26 runs were performed using two different nozzle configurations with pressures in the test capsule around 0.3 psia. No problems were encountered during the test series resulting from the optical hardware being located in the test capsule and acceptable resolution images were captured. The test campaign demonstrated the ability of using the BOS technique for small, rocket engine, plume flow visualization during altitude testing.