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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 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1994

Acoustics of laminar boundary layers breakdown

Boundary layer flow transition has long been suggested as a potential noise source in both marine (sonar-dome self noise) and aeronautical (aircraft cabin noise) applications, owing to the highly transient nature of process. The design of effective noise control strategies relies upon a clear understanding of the source mechanisms associated with the unsteady flow dynamics during transition. Due to formidable mathematical difficulties, theoretical predictions either are limited to early linear and weakly nonlinear stages of transition, or employ acoustic analogy theories based on approximate source field data, often in the form of empirical correlation. In the present work, an approach which combines direct numerical simulation of the source field with the Lighthill acoustic analogy is utilized. This approach takes advantage of the recent advancement in computational capabilities to obtain detailed information about the flow-induced acoustic sources. The transitional boundary layer flow is computed by solving the incompressible Navier-Stokes equations without model assumptions, thus allowing a direct evaluation of the pseudosound as well as source functions, including the Lighthill stress tensor and the wall shear stress. The latter are used for calculating the radiated pressure field based on the Curle-Powell solution of the Lighthill equation. This procedure allows a quantitative assessment of noise source mechanisms and the associated radiation characteristics during transition from primary instability up to the laminar breakdown stage. In particular, one is interested in comparing the roles played by the fluctuating volume Reynolds stress and the wall-shear-stresses, and in identifying specific flow processes and structures that are effective noise generators.

Wang, Meng↗
NASA NTRS1981

Aerobraking techniques for planetary missions

Aerobraking techniques are applied to trajectory design for the VOIR (Venus Orbiting Imaging Radar) mission to enhance overall mission performance. The approach can be used for other planetary missions such as a Titan orbiter or a Mars sample return to earth. The orientation of the elliptical insertion orbit must be chosen in such a way that the combined effects of solar gravity, central-body harmonics, and drag on periapsis altitude are minimized to avoid frequent maneuvers. The selection of the base and width of a corridor for periapsis altitude depends upon the temperature and integrated heat of the aerobrake shield, the atmospheric density profile, the duration of aerobraking, the number of maneuvers, the time available between maneuvers, and uncertainties in the central-body gravity field and atmospheric density. Flight path profiles for an aerodynamically stable spacecraft and an inertially fixed spacecraft passing through the free molecular flow regime, as well as the transitional flow regime, are illustrated

Tang, C. C. H.↗
NASA NTRS1993

Three-dimensional hypersonic rarefied flow calculations using direct simulation Monte Carlo method

A summary of three-dimensional simulations on the hypersonic rarefied flows in an effort to understand the highly nonequilibrium flows about space vehicles entering the Earth's atmosphere for a realistic estimation of the aerothermal loads is presented. Calculations are performed using the direct simulation Monte Carlo method with a five-species reacting gas model, which accounts for rotational and vibrational internal energies. Results are obtained for the external flows about various bodies in the transitional flow regime. For the cases considered, convective heating, flowfield structure and overall aerodynamic coefficients are presented and comparisons are made with the available experimental data. The agreement between the calculated and measured results are very good.

Celenligil, M. Cevdet↗
NASA NTRS2003

Computational And Experimental Studies Of Three-Dimensional Flame Spread Over Liquid Fuel Pools

Schiller, Ross, and Sirignano (1996) studied ignition and flame spread above liquid fuels initially below the flashpoint temperature by using a two-dimensional computational fluid dynamics code that solves the coupled equations of both the gas and the liquid phases. Pulsating flame spread was attributed to the establishment of a gas-phase recirculation cell that forms just ahead of the flame leading edge because of the opposing effect of buoyancy-driven flow in the gas phase and the thermocapillary-driven flow in the liquid phase. Schiller and Sirignano (1996) extended the same study to include flame spread with forced opposed flow in the gas phase. A transitional flow velocity was found above which an originally uniform spreading flame pulsates. The same type of gas-phase recirculation cell caused by the combination of forced opposed flow, buoyancy-driven flow, and thermocapillary-driven concurrent flow was responsible for the pulsating flame spread. Ross and Miller (1998) and Miller and Ross (1998) performed experimental work that corroborates the computational findings of Schiller, Ross, and Sirignano (1996) and Schiller and Sirignano (1996). Cai, Liu, and Sirignano (2002) developed a more comprehensive three-dimensional model and computer code for the flame spread problem. Many improvements in modeling and numerical algorithms were incorporated in the three-dimensional model. Pools of finite width and length were studied in air channels of prescribed height and width. Significant three-dimensional effects around and along the pool edge were observed. The same three-dimensional code is used to study the detailed effects of pool depth, pool width, opposed air flow velocity, and different levels of air oxygen concentration (Cai, Liu, and Sirignano, 2003). Significant three-dimensional effects showing an unsteady wavy flame front for cases of wide pool width are found for the first time in computation, after being noted previously by experimental observers (Ross and Miller, 1999). Regions of uniform and pulsating flame spread are mapped for the flow conditions of pool depth, opposed flow velocity, initial pool temperature, and air oxygen concentration under both normal and microgravity conditions. Details can be found in Cai et al. (2002, 2003). Experimental results recently performed at NASA Glenn of flame spread across a wide, shallow pool as a function of liquid temperature are also presented here.

Ross, Howard D.↗
NASA NTRSDate not supplied

Fluid-Thermal-Structural Interactions Induced by an Asymmetric Shock-Wave/Boundary-Layer Interaction in a Mach-6 Compression Corner

An experimental study is conducted of the fluid-thermal-structural interaction of a clamped compliant panel exposed to a three dimensional shock-wave/boundary-layer interaction (SWBLI) induced by a Mach-6 compression ramp with a spanwise nonuniform incoming boundary layer. The nonuniform boundary layer was produced by placing trips on one side of the upstream flat plate, resulting in largely turbulent flow on the tripped side and transitional flow on the untripped side. Measurements of the flowfield confirmed that the tripped boundary layer contained elevated levels of unsteadiness, and the SWBLI was observed to vary from attached to fully separated as the ramp angle was increased from 10◦ to 38◦; the separation region on the tripped side of the panel was noticeably smaller, showing the elevated turbulence levels of the tripped-side flow to remain relatively localized rather than diffusing across the whole model. Full-field, time-resolved panel deformations were measured using high-speed photogrammetry and the vibrational response at each compression angle was characterized. Although the measured modes conformed largely to those from classical clamped-plate theory, some skewing of the mode shapes was observed. IR thermography highlighted regions of the compliant region where elevated temperatures were likely to promote thermal softening effects to the transient panel response. The quasi-static deformation and stress field was used to characterize the internal stress factor of each mode and showed a meaningful relationship between transient panel response and stress contained within each mode: modes with antinodes lying in high-stress areas of the plate tended to exhibit increases in vibrational frequency and decreases in vibrational power, whereas the opposite was true for modes with antinodes in low-stress areas.

Spectral Proper Orthogonal Decomposition↗
NASA NTRS2005

3-D High-Lift Flow-Physics Experiment - Transition Measurements

An analysis of the flow state on a trapezoidal wing model from the NASA 3-D High Lift Flow Physics Experiment is presented. The objective of the experiment was to characterize the flow over a non-proprietary semi-span three-element high-lift configuration to aid in assessing the state of the art in the computation of three-dimensional high-lift flows. Surface pressures and hot-film sensors are used to determine the flow conditions on the slat, main, and flap. The locations of the attachments lines and the values of the attachment line Reynolds number are estimated based on the model surface pressures. Data from the hot-films are used to determine if the flow is laminar, transitional, or turbulent by examining the hot-film time histories, statistics, and frequency spectra.

McGinley, Catherine B.↗
NASA NTRS2007

Ignition Characterization Test Results for the LO2/Ethanol Propellant Combination

A series of contracts were issued by the Marshall Space Flight Center (MSFC) of the National Aeronautics and Space Administration (NASA) un der the auspices of the Exploration Systems Mission Directorate to de velop and expand the maturity of candidate technologies considered to be important for future space exploration. One such technology was to determine the viability of incorporating non-toxic propellants for R eaction Control Subsystems (RCS). Contract NAS8-01109 was issued to A erojet to develop a dual thrust Reaction Control Engine (RCE) that ut ilized liquid oxygen and ethanol as the propellants. The dual thrust RCE incorporated a primary thrust level of 870 lbf, and a vernier thru st level of 10 - 30 lbf. The preferred RCS approach for the dual thru st RCE was to utilize pressure-fed liquid oxygen (LOX) and ethanol pr opellants; however, previous dual thrust feasibility testing incorporated GOX/Ethanol igniters as opposed to LOX/Ethanol igniters in the de sign. GOX/Ethanol was easier to ignite, but this combination had syst em design implications of providing GOX for the igniters. A LOX/Ethan ol igniter was desired; however, extensive LOX/Ethanol ignition data over the anticipated operating range for the dual thrust RCE did not e xist. Therefore, Aerojet designed and tested a workhorse LOX igniter to determine LOX/Ethanol ignition characteristics as part of a risk m itigation effort for the dual thrust RCE design. The objective of the ignition testing was to demonstrate successful ignition from GOX to LOX, encompassing potential two-phase flow conditions anticipated being present in real mission applications. A workhorse igniter was desig ned to accommodate the full LOX design flowrate, as well as a reduced GOX flowrate. It was reasoned that the initial LOX flow through the igniter would flash to GOX due to the latent heat stored in the hardwa re, causing a reduced oxygen flowrate because of a choked, or sonic, flow condition through the injection elements. As LOX flow continued, the hardware would chill-in, with the injected oxygen flow transitioning from cold GOX through two'phase flow to subcooled LOX. The Workh orse igniter was well instrumented: Pressure and temperature instrumentation permitted oxygen state points to be determined in the igniter oxidizer manifold, and gas-side igniter chamber thermocouples provide d chamber thermal profile characteristics. The cold flow chamber pres sure (Pc) for each test was determined and coupled with the igniter chamber diameter (De) to calculate the characteristic quench parameter (Pc x Dc), which was plotted as a function of core mixture ratio, MRc . Ignition limits were determined over a broad range of valve inlet conditions, and ignition was demonstrated with oxygen inlet conditions that ranged from subcooled 210 deg R LOX to 486 deg R GOX. Once ign ited at cold GOX conditions, combustion was continuous as the hardwar e chilled in and the core mixture ratio transitioned from values near 1.0 to over 12.5. Pulsing is required in typical RCS engines; therefore, the workhorse igniter was pulse tested to verify the ability to pr ovide the required ignition for a pulsing RCE. The minimum electrical pulse width (EPW) of the dual thrust RCE was 0.080 seconds. Igniter pulse tests were performed at three conditions: (1) an EPW of 0.080 se conds at 25% duty cycle for 400 pulses; (2) an EPW of 0.160 seconds a nd a 5% duty cycle for 124 pulses; (3) an EPW of 0.160 seconds and a 50% duty cycle for 380 pulses. Successful ignition of LOX/Ethanol was demonstrated over a broad range of valve inlet conditions, with the empirically determined LOX/Ethanol ignition limits extending the previous database established for GOX/Ethanol ignition limits. Although th e observed chill-in characteristics of the hardware varied significan tly with flowrate, ignition was readily achieved. Combustion was marg inal at extremely fuel-rich conditions, and it fluctuated as the oxygen passed rough the twophase flow regime during the period of hardware chill-in. Pulse testing showed good repeatability with 100 percent r e-ignition for all pulses. Certain pulse-to-pulse repeatability requirements for actual RCS operation may necessitate establishment of mini mum oxygen flow rates and engine thrust levels for satisfactory engin e performance.

Robinson, Philip J.↗
NASA NTRS1987

Inflectional instabilities in the wall region of bounded turbulent shear flows

The primary thrust of this research was to identify one or more mechanisms responsible for strong turbulence production events in the wall region of bounded turbulent shear flows. Based upon previous work in a transitional boundary layer, it seemed highly probable that the production events were preceded by an inflectional velocity profile which formed on the interface between the low-speed streak and the surrounding fluid. In bounded transitional flows, this unstable profile developed velocity fluctuations in the streamwise direction and in the direction perpendicular to the sheared surface. The rapid growth of these instabilities leads to a breakdown and production of turbulence. Since bounded turbulent flows have many of the same characteristics, they may also experience a similar type of breakdown and turbulence production mechanism.

Swearingen, Jerry D.↗
NASA NTRS2001

Normal Gravity Testing of a Microchannel Phase Separator for In Situ Resource Utilization

A microchannel separator, with 2.7 millimeters as the smallest dimension, was tested, and a pore throat structure captured and removed liquid from a gas-liquid stream. The microchannel device was tested over a of gas and liquid flow rates ranging from 0.0005 up to 0. 14 volume fraction of liquid. Four liquids were tested with air. The biggest factor affecting the throughput is the capacity of liquid flow through the pore throat, which is dictated by permeability, liquid viscosity, flow area, pore throat thickness, and pressure difference across the pore throat. Typically, complete separation of gas and liquid fractions was lost when the liquid flow rate reached about 40 to 60% of the pore throat capacity. However, this could occur over a range of 10 to 90% utilization of pore throat capacity. Breakthrough occurs in the microchannel phase separator at conditions similar to the annular to plug flow transition of two-phase microgravity pipe flow implying that operating in the proper flow regime is crucial. Analysis indicates that the Bond number did not affect performance, supporting the premise that hydrodynamic, interfacial, and capillary forces are more important than gravity. However, the relative importance of gravity is better discerned through testing under reduced gravity conditions.

TeGrotenhuis, Ward E.↗
NASA NTRS1986

Shuttle Tethered Aerothermodynamics Research Facility (STARFAC)

The Earth's atmosphere from 90 km to 200 km provides the last aerothermodynamics frontier. Present NASA programs which require but also can provide an understanding of the aerodynamics and aerothermodynamics of the free molecule and transition flows that exist at these altitudes are the Aeroassisted OTV, Entry Research Vehicle and the Tethered Satellite. Each of these programs provides a unique opportunity to do flight research in the rarefied upper atmosphere. However, the Tethered Satellite Program provides, because of its capability to obtain global, in-situ, steady state data, the greatest potential to: (1)define the performance of aerodynamic shapes as a function of environmental characteristics (free molecule, transition, slip flow regimes); (2)define the characteristics of the upper atmosphere and the global variability of properties such as composition temperature, pressure and density. Such data are required to accomplish the systematic development and verification of analytical prediction techniques required to support advance configuration designs.

Siemers, P. M.↗
NASA NTRS1993

Experimental cavity pressure measurements at subsonic and transonic speeds. Static-pressure results

An experimental investigation was conducted to determine cavity flow-characteristics at subsonic and transonic speeds. A rectangular box cavity was tested in the Langley 8-Foot Transonic Pressure Tunnel at Mach numbers from 0.20 to 0.95 at a unit Reynolds number of approximately 3 x 10(exp 6) per foot. The boundary layer approaching the cavity was turbulent. Cavities were tested over a range of length-to-depth ratios (l/h) of 1 to 17.5 for cavity width-to-depth ratios of 1, 4, 8, and 16. Fluctuating- and static-pressure data in the cavity were obtained; however, only static-pressure data is analyzed. The boundaries between the flow regimes based on cavity length-to-depth ratio were determined. The change to transitional flow from open flow occurs at l/h at approximately 6-8 however, the change from transitional- to closed-cavity flow occurred over a wide range of l/h and was dependent on Mach number and cavity configuration. The change from closed to open flow as found to occur gradually. The effect of changing cavity dimensions showed that if the vlaue of l/h was kept fixed but the cavity width was decreased or cavity height was increased, the cavity pressure distribution tended more toward a more closed flow distribution.

Plentovich, E. B.↗
NASA NTRS2000

Flow Simulation of Solid Rocket Motors: Sub-Scale Air Flow Simulation of Port Flows - 2

The injection-flow issuing from a porous medium in the cold-flow simulation of internal port flows in solid rocket motors is characterized by a spatial instability termed pseudoturbulence that produces a rather non-uniform (lumpy) injection-velocity profile. The objective of this study is to investigate the interaction between the injection- and the developing axial-flows. The findings show that this interaction generally weakens the lumpy injection profile and affects the subsequent development of the axial flow. The injection profile is found to depend on the material characteristics, and the ensuing pseudoturbulence is a function of the injection velocity, the axial position and the distance from the porous wall. The flow transition (from laminar to turbulent) of the axial-flow is accelerated in flows emerging from smaller pores primarily due to the higher pseudoturbulence produced by the smaller pores in comparison to that associated with larger pores. In flows with rather uniform injection-flow profiles (weak or no pseudoturbulence), the axial and transverse velocity components in the porous duct are found to satisfy the sine/cosine analytical solutions derived from inviscid assumptions. The transition results from the present study are compared with previous results from surveyed literature, and detailed flow development measurements are presented in terms of the blowing fraction, and characterizing Reynolds numbers.

Yeh, Y. P.↗
NASA NTRS1996

Hypersonic Boundary Layer Stability over a Flared Cone in a Quiet Tunnel

Hypersonic boundary layer measurements were conducted over a flared cone in a quiet wind tunnel. The flared cone was tested at a freestream unit Reynolds number of 2.82x106/ft in a Mach 6 flow. This Reynolds number provided laminar-to-transitional flow over the model in a low-disturbance environment. Point measurements with a single hot wire using a novel constant voltage anemometry system were used to measure the boundary layer disturbances. Surface temperature and schlieren measurements were also conducted to characterize the laminar-to-transitional state of the boundary layer and to identify instability modes. Results suggest that the second mode disturbances were the most unstable and scaled with the boundary layer thickness. The integrated growth rates of the second mode compared well with linear stability theory in the linear stability regime. The second mode is responsible for transition onset despite the existence of a second mode sub-harmonic. The sub-harmonic wavelength also scales with the boundary layer thickness. Furthermore, the existence of higher harmonics of the fundamental suggests that non-linear disturbances are not associated with high free stream disturbance levels.

Lachowicz, Jason T.↗
NASA NTRS2006

Ignition Characterization Test Results for the LO2/Ethanol Propellant Combination

A series of contracts were issued by the Marshall Space Flight Center (MSFC) of the National Aeronautics and Space Administration (NASA) under the auspices of the Exploration Systems Mission Directorate to develop and expand the maturity of candidate technologies considered to be important for future space exploration. One such technology was to determine the viability of incorporating non-toxic propellants for Reaction Control Subsystems (RCS). Contract NAS8-01109 was issued to Aerojet to develop a dual thrust Reaction Control Engine (RCE) that utilized liquid oxygen and ethanol as the propellants. The dual thrust RCE incorporated a primary thrust level of 870 lbf, and a vernier thrust level of 10 - 30 lbf. The preferred RCS approach for the dual thrust RCE was to utilize pressure-fed liquid oxygen (LOX) and ethanol propellants; however, previous dual thrust feasibility testing incorporated GOX/Ethanol igniters as opposed to LOX/Ethanol igniters in the design. GOX/Ethanol was easier to ignite, but this combination had system design implications of providing GOX for the igniters. A LOX/Ethanol igniter was desired; however, extensive LOX/Ethanol ignition data over the anticipated operating range for the dual thrust RCE did not exist. Therefore, Aerojet designed and tested a workhorse LOX igniter to determine LOX/Ethanol ignition characteristics as part of a risk mitigation effort for the dual thrust RCE design. LOX, encompassing potential two-phase flow conditions anticipated being present in real mission applications. A workhorse igniter was designed to accommodate the hll LOX design flowrate, as well as a reduced GOX flowrate. It was reasoned that the initial LOX flow through the igniter would flash to GOX due to the latent heat stored in the hardware, causing a reduced oxygen flowrate because of a choked, or sonic, flow condition through the injection elements. As LOX flow continued, the hardware would chill-in, with the injected oxygen flow transitioning from cold GOX through two-phase flow to subcooled LOX. permitted oxygen state points to be determined in the igniter oxidizer manifold, and gas-side igniter chamber thermocouples provided chamber thermal profile characteristics. The cold flow chamber pressure (P(sub c)) for each test was determined and coupled with the igniter chamber diameter (D(sub c)) to calculate the characteristic quench parameter (P(sub c) x D(sub c)), which was plotted as a function of core mixture ratio, m. Ignition limits were determined over a broad range of valve inlet conditions, and ignition was demonstrated with oxygen inlet conditions that ranged from subcooled 210 R LOX to 486 R GOX. Once ignited at cold GOX conditions, combustion was continuous as the hardware chilled in and the core mixture ratio transitioned from values near 1.0 to over 12.5. Pulsing is required in typical RCS engines; therefore, the workhorse igniter was pulse tested to verify the ability to provide the required ignition for a pulsing RCE. The minimum electrical pulse width (EPW) of the dual thrust RCE was 0.080 seconds.

Popp, Christopher G.↗
NASA NTRS2004

Ignition Characterization Tests of the LOX/Ethanol Propellant Combination

A series of contracts have been issued by the Marshall Space Flight Center (MSFC) of the National Aeronautics and Space Administration (NASA) to explore candidate technologies considered to be important for the Next Generation Launch Technology (NGLT) effort. One aspect of the NGLT effort is to explore the potential of incorporating non-toxic propellants for Reaction Control Subsystems (RCS). Contract NAS8-01109 has been issued to Aerojet to develop a dual thrust Reaction Control Engine (RCE) that utilizes liquid oxygen and ethanol as the propellants. The dual thrust RCE incorporates a primary thrust level of 870 lbf, and a vernier thrust level of 10 - 30 lbf. Aerojet has designed and tested a workhorse LOX igniter to determine LOX/Ethanol ignition characteristics as part of a risk mitigation effort for the dual thrust RCE design. The objective of the ignition testing was to demonstrate successfid ignition from GOX to LOX, encompassing potential two-phase flow conditions. The workhorse igniter was designed to accommodate the full LOX design flowrate, as well as a reduced GOX flowrate. It was reasoned that the initial LOX flow through the igniter would flash to GOX due to the inherent heat stored in the hardware, causing a reduced oxygen flowrate because of a choked, or sonic, flow condition through the injection elements. As LOX flow continued, the inherent heat of the test hardware would be removed and the hardware would chill-in, with the injected oxygen flow transitioning from cold GOX through two-phase flow to subcooled LOX. Pressure and temperature instrumentation permitted oxygen state points to be determined, and gas-side igniter chamber thermocouples provided chamber thermal profile characteristics. The cold flow chamber pressure (P(sub c)) for each test was determined and coupled with the igniter chamber diameter (D(sub c)) to calculate the characteristic quench parameter (P(sub c) x D(sub c)), which was plotted as a function of core mixture ratio, MR(sub c). Ignition limits were determined over a broad range of valve inlet conditions, and ignition was demonstrated with oxygen inlet conditions that ranged from subcooled 173 R LOX to 480 R GQX. Once ignited at cold GOX conditions, combustion was continuous as the hardware chilled in and the core mixture ratio transitioned from values near 1.0 to over 12.5.

Popp, Christopher G.↗
NASA NTRS1987

Details of the computed flowfield over a circular cylinder at Reynolds number 1200

The application of an upwind-biased implicit approximate factorization Navier-Stokes algorithm to the unsteady impulsive start-up flow over a circular cylinder at Reynolds number 1200 is described. The complete form of the compressible Navier-Stokes equations is used, and the algorithm is second-order accurate in both space and time. The development with time of the shape and size of the separated vortical flow region is computed, as well as the time-variation of several boundary layer parameters and profile shapes. Computations, in general, show excellent agreement with experiment, although the present method predicts a more rapid onset of reversed flow on the cylinder than evidenced in experiment. The changes that the vortical region behind the cylinder undergoes as the symmetric flow transitions to periodic vortex shedding are discussed. The flow becomes periodic with a Strouhal frequency of 0.222, which compares well with the experimental value of approximately 0.21. The effect of grid density on the development of the unsteady flow is also shown.

Rumsey, C. L.↗