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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 253 records · Page 14
All sourcesNASA NTRSDOE OSTI
NASA NTRS2001

The FU Orionis Phenomenon and Solar Nebula Material

We summarize astronomical, meteoritic, and theoretical evidence relating to the FU Orionis phenomenon. This evidence suggests that at early times (the first few 10(exp 5) yr), the solar nebula experienced a hot phase characterized by high accretion rates (the "FU Ori epoch"), punctuated by episodic outbursts of enhanced mass flow through the inner part of the disk (less than or equal to 0.3 AU). Throughout this epoch, disk midplane temperatures exceeded 1000 K at 1 AU. Diminishing infall from the cloud core led to decreasing mass flux throughout the disk. When mass flow de creased below the value critical for outburst (5 x 10(exp -7) solar mass/yr, as suggested by thermal ionization instability models), outbursts ceased and the T Tauri epoch began. Outburst timescales are too long to explain calcium- and aluminum-rich inclusion (CAI) and chondrule formation. Volatility-dependent fractionation patterns seen in meteoritic materials suggest that solids formed beginning during a hot epoch when temperatures exceeded 1400 K, and the presence of volatiles in chondrites argues that this process continued until the nebula had cooled to below 400 K. The thermal ionization instability model for FU Ori outbursts is in quantitative agreement with astronomical observations. Its results imply that the terrestrial region of the nebula reached the hot end of this range only during a time when mass flow through the disk was high enough to trigger outbursts (i.e., the FU Ori epoch) and reached the cool end of this range only during the later T Tauri epoch. According to the models, heating of material in the terrestrial planet region during individual FU Ori outbursts would be limited to surface layers of the nebula, leaving midplane materials (which are at greater than or equal to 1000 K) largely unaffected. Alternative FU Ori models should be developed, particularly if compositional differences among chondrite clans are attributable to episodic heating.

Bell, K. R.↗
NASA NTRSDate not supplied

Model Preparation Areas for Propulsion Airframe Integration Testing at the NASA Langley Research Center

Currently, there are three major NASA Langley facilities that perform Propulsion Airframe Integration (PAI) testing. These include the National Transonic Facility (NTF), the 14- by 22-Foot Subsonic Tunnel, and the 4-Foot Supersonic Unitary Plan Wind Tunnel. While each of these facilities have their respective model preparation areas, one model preparation area, located at the NTF complex, focuses on high mass flow blowing systems in support of other Langley Research Center facilities. The application of PAI requires special attention to the air delivery system interaction with the model. These interactions focus on the tares across the force and moment balance. The interactions are a result of the mass flow and pressure of the air through the system and the efficiency of the inlet and nozzle geometries. While this paper does not characterize any specific propulsion system, it does focus on the mass flow and pressure requirements for risk-reduction planning of such systems prior to installation in a designated wind tunnel. An example of model and nozzle sizing will be presented for the Tail Cone Thruster (TCT) variant of the Common Research Model (CRM).

PAI↗
NASA NTRS1972

Mass transport contamination study

A theoretical analysis was performed to determine the effects of outgassing and waste dumping on the contamination field around an orbiting spacecraft. The spacecraft was assumed to be spherical in shape with the mass flow emitting uniformly from the spherical surface at a constant rate and in a D'Lambertian spatial distribution. The outflow of gases were assumed to be neutrally charged and of a single species with a molecular weight characteristic of a composite of the actual species involved in the mass flow. The theoretical analysis showed that, for outgassing only, less than 1.5 percent of the outgas products will return to the Skylab spacecraft as a result of intermolecular collisions. When the total mass flow from the spacecraft, including waste dumps and reaction control motor firings, was considered, it was estimated that about 30 percent will return to the spacecraft.

Robertson, S. J.↗
NASA NTRS2008

Boundary-Layer-Ingesting Inlet Flow Control

An experimental study was conducted to provide the first demonstration of an active flow control system for a flush-mounted inlet with significant boundary-layer-ingestion in transonic flow conditions. The effectiveness of the flow control in reducing the circumferential distortion at the engine fan-face location was assessed using a 2.5%-scale model of a boundary-layer-ingesting offset diffusing inlet. The inlet was flush mounted to the tunnel wall and ingested a large boundary layer with a boundary-layer-to-inlet height ratio of 35%. Different jet distribution patterns and jet mass flow rates were used in the inlet to control distortion. A vane configuration was also tested. Finally a hybrid vane/jet configuration was tested leveraging strengths of both types of devices. Measurements were made of the onset boundary layer, the duct surface static pressures, and the mass flow rates through the duct and the flow control actuators. The distortion and pressure recovery were measured at the aerodynamic interface plane. The data show that control jets and vanes reduce circumferential distortion to acceptable levels. The point-design vane configuration produced higher distortion levels at off-design settings. The hybrid vane/jet flow control configuration reduced the off-design distortion levels to acceptable ones and used less than 0.5% of the inlet mass flow to supply the jets.

Owens, Lewis R.↗
NASA NTRS1995

Cooling Panel Optimization for the Active Cooling System of a Hypersonic Aircraft

Optimization of cooling panels for an active cooling system of a hypersonic aircraft is explored. The flow passages are of rectangular cross section with one wall heated. An analytical fin-type model for incompressible flow in smooth-wall rectangular ducts with coupled wall conduction is proposed. Based on this model, the a flow rate of coolant to each design minimum mass flow rate or coolant for a single cooling panel is obtained by satisfying hydrodynamic, thermal, and Mach number constraints. Also, the sensitivity of the optimal mass flow rate of coolant to each design variable is investigated. In addition, numerical solutions for constant property flow in rectangular ducts, with one side rib-roughened and coupled wall conduction, are obtained using a k-epsilon and wall function turbulence model, these results are compared with predictions of the analytical model.

Youn, B.↗
NASA NTRS1977

Experimental evaluation of nacelle-airframe interference forces and pressures at Mach numbers of 0.9 to 1.4

Detailed interference force-and-pressure data were obtained on a representative supersonic transport wing-body-nacelle combination at Mach numbers of 0.9 to 1.4. The basic model consisted of a delta wing-body aerodynamic model with a length of 158.0 cm (62.2 in.) and a wingspan of 103.6 cm (40.8 in.) and four independently supported nacelles positioned beneath the model. The experimental program was conducted in the Ames 11- by 11-Foot Wind Tunnel at a constant unit Reynolds number. The primary variables examined included Mach number, angle of attack, nacelle position, and nacelle mass-flow ratio. Under the most favorable conditions, the net interference drag was equal to 50 percent the drag of four isolated nacelles at M = 1.4, 75 percent at M = 1.15, and 144 percent at M = 0.90. The overall interference effects were found to be rather constant over the operating angle-of-attack range of the configuration. The effects of mass-flow ratio on the interference pressure distributions were limited to the lip region of the nacelle and the local wing surface in the immediate vicinity of the nacelle lip. The net change in the measured interference forces resulting from variations in the nacelle mass-flow ratio were found to be quite small.

Bencze, D. P.↗
NASA NTRS1995

Jet Mixing and Emission Characteristics of Transverse Jets in Annular and Cylindrical Confined Crossflow

Three dimensional turbulent reacting CFD analyses were performed on transverse jets injected into annular and cylindrical (can) confined crossflows. The goal was to identify and assess mixing differences between annular and can geometries. The approach taken was to optimize both annular and can configurations by systematically varying orifice spacing until lowest emissions were achieved, and then compare the results. Numerical test conditions consisted of a jet-to-mainstream mass-flow ratio of 3.2 and a jet-to-mainstream momentum-flux ratio (J) of 30. The computational results showed that the optimized geometries had similar emission levels at the exit of the mixing section although the annular configuration did mix-out faster. For lowest emissions, the density correlation parameter (C = (S/H) square root of J) was 2.35 for the annular geometry and 3.5 for the can geometry. For the annular geometry, the constant was about twice the value seen for jet mixing at low mass-flow ratios (i.e., MR less than 0.5). For the can geometry, the constant was about 1 1/2 times the value seen for low mass-flow ratios.

Bain, D. B.↗
NASA NTRS1951

Low-Lift Drag and Duct Pressure Recovery of a 1/8.25-Scale Model of the Consolidated Vultee XF-92 Airplane at Mach Numbers from 0.7 to 1.4

A flight investigation has been made to determine the external drag and pressure recovery of a 1/8.25 - scale flight model of the Consolidated Vultee XF-92 from Mach numbers 0.7 to 1.4 and Reynolds numbers from 8.5 x 10(exp 6) to 19.2 x 10(exp 6) at or near zero lift. Relative mass flow, average pressure recovery, total drag, internal drag, and external drag are presented as functions of Mach number. Between Mach numbers of 0.90 and 0.975, the external drag of the configuration (including base drag of the inner body and additive drag) was about equal to that of a similar model with a faired nose and no mass flow; however, at supersonic speeds the drag coefficient for the faired-nose model remained relatively constant whereas the drag coefficient for the ducted model continued to increase sharply. The internal drag coefficient of the duct was roughly constant at 0.013 up to a Mach number of 1.20; after which it decreased to 0.0075 at a Mach number of 1.4. The over-all pressure recovery of the inlet and duct varied from 94 percent at a Mach number of 0.7 to about 91 percent at a Mach number of 1.4 at a relative-mass-flow ratio of about 0.30. The losses in pressure recovery were believed to be caused by the possible occurrence of separation of flow from the inner body and by an aerodynamically unclean internal configuration which did not duplicate the form proposed for the original XF-92 airplane.

Mitcham, Grady L.↗
NASA NTRS1995

GHRS observations of mass-loaded flows in Abell 78

Spectroscopic observations of the central star of the planetary nebula Abell 78 were obtained with the Goddard High Resolution Spectrograph (GHRS) onboard the Hubble Space Telescope (HST) in the vicinity of the C IV lambda 1548.2, 1550.8 doublet. We find a series of narrow absorption features superposed on the broad, P Cygni stellar wind profile. These features are seen in both components of the doublet at heliocentric radial velocities of -18, -71, -131, and -192 km/s. At higher velocities, individual components are no longer distinct but, rather, merge into a continuous absorption extending to approximately -385 km/s. This is among the highest velocities ever detected for gas in a planetary nebula. The -18 km/s feature originates in an outer envelope of normal composition, while the -71 km/s feature is produced in the wind-swept shell encircling an irregular wind-blown bubble in the planetary nebula center. The hydrogen-poor ejecta of Abell 78, consisting of dense knots with wind-blown tails, are located in the bubble's interior, in the vicinity of the stellar wind termination shock. The high-velocity C IV lambda 154 absorption features can be explained as due to parcels of ejecta being accelerated to high velocities as they are swept up by the stellar wind during its interaction with dense condensations of H-poor ejecta. As the ablated material is accelerated, it will partially mix with the stellar wind, creating a mass-loaded flow. The abundance anomalies seen at the rim of the bubble attest to the transport of H-poor knot material by such a flow.

Harrington, J. Patrick↗
NASA NTRS2015

Loop Heat Pipe Temperature Oscillation Induced by Gravity Assist and Reservoir Heating

The Laser Thermal Control System (LCTS) for the Advanced Topographic Laser Altimeter System (ATLAS) to be installed on NASA's Ice, Cloud, and Land Elevation Satellite (ICESat-2) consists of a constant conductance heat pipe and a loop heat pipe (LHP) with an associated radiator. During the recent thermal vacuum testing of the LTCS where the LHP condenser/radiator was placed in a vertical position above the evaporator and reservoir, it was found that the LHP reservoir control heater power requirement was much higher than the analytical model had predicted. Even with the control heater turned on continuously at its full power, the reservoir could not be maintained at its desired set point temperature. An investigation of the LHP behaviors found that the root cause of the problem was fluid flow and reservoir temperature oscillations, which led to persistent alternate forward and reversed flow along the liquid line and an imbalance between the vapor mass flow rate in the vapor line and liquid mass flow rate in the liquid line. The flow and temperature oscillations were caused by an interaction between gravity and reservoir heating, and were exacerbated by the large thermal mass of the instrument simulator which modulated the net heat load to the evaporator, and the vertical radiator/condenser which induced a variable gravitational pressure head. Furthermore, causes and effects of the contributing factors to flow and temperature oscillations intermingled.

Thermal Control Systems↗
NASA NTRS2015

Loop Heat Pipe Temperature Oscillation Induced by Gravity Assist and Reservoir Heating

The Laser Thermal Control System (LCTS) for the Advanced Topographic Laser Altimeter System (ATLAS) to be installed on NASA's Ice, Cloud, and Land Elevation Satellite (ICESat-2) consists of a constant conductance heat pipe and a loop heat pipe (LHP) with an associated radiator. During the recent thermal vacuum testing of the LTCS where the LHP condenser/radiator was placed in a vertical position above the evaporator and reservoir, it was found that the LHP reservoir control heater power requirement was much higher than the analytical model had predicted. Even with the control heater turned on continuously at its full power, the reservoir could not be maintained at its desired set point temperature. An investigation of the LHP behaviors found that the root cause of the problem was fluid flow and reservoir temperature oscillations, which led to persistent alternate forward and reversed flow along the liquid line and an imbalance between the vapor mass flow rate in the vapor line and liquid mass flow rate in the liquid line. The flow and temperature oscillations were caused by an interaction between gravity and reservoir heating, and were exacerbated by the large thermal mass of the instrument simulator which modulated the net heat load to the evaporator, and the vertical radiator/condenser which induced a variable gravitational pressure head. Furthermore, causes and effects of the contributing factors to flow and temperature oscillations intermingled.

0000↗
NASA NTRS2004

Turbomachinery for Low-to-High Mach Number Flight

The thrust capability of turbojet cycles is reduced at high flight Mach number (3+) by the increase in inlet stagnation temperature. The 'hot section' temperature limit imposed by materials technology sets the maximum heat addition and, hence, sets the maximum flight Mach number of the operating envelope. Compressor pre-cooling, either via a heat exchanger or mass-injection, has been suggested as a means to reduce compressor inlet temperature and increase mass flow capability, thereby increasing thrust. To date, however, no research has looked at compressor cooling (i.e., using a compressor both to perform work on the gas path air and extract heat from it simultaneously). We wish to assess the feasibility of this novel concept for use in low-to-high Mach number flight. The results to-date show that an axial compressor with cooling: (1) relieves choking in rear stages (hence opening up operability), (2) yields higher-pressure ratio and (3) yields higher efficiency for a given corrected speed and mass flow. The performance benefit is driven: (i) at the blade passage level, by a decrease in the total pressure reduction coefficient and an increase in the flow turning; and (ii) by the reduction in temperature that results in less work required for a given pressure ratio. The latter is a thermodynamic effect. As an example, calculations were performed for an eight-stage compressor with an adiabatic design pressure ratio of 5. By defining non-dimensional cooling as the percentage of compressor inlet stagnation enthalpy removed by a heat sink, the model shows that a non-dimensional cooling of percent in each blade row of the first two stages can increase the compressor pressure ratio by as much as 10-20 percent. Maximum corrected mass flow at a given corrected speed may increase by as much as 5 percent. In addition, efficiency may increase by as much as 5 points. A framework for characterizing and generating the performance map for a cooled compressor has been developed. The approach is based upon CFD computations and mean line analysis. Figures of merit that characterize the bulk performance of blade passage flows with and without cooling are extracted from CFD solutions. Such performance characterization is then applied to a preliminary compressor design framework (mean line). The generic nature of this approach makes it suitable for assessing the effect of different types of compressor cooling schemes, such as heat exchange or evaporative cooling (mass injection). Future work will focus on answering system level questions regarding the feasibility of compressor cooling. Specifically, we wish to determine the operational parametric space in which compressor cooling would be advantageous over other high flight Mach number propulsion concepts. In addition, we will explore the design requirements of cooled compressor turbomachinery, as well as the flow phenomena that limit and control its operation, and the technology barriers that must be crossed for its implementation.

Tan, Choon S.↗
NASA NTRS2009

Liquid Oxygen/Liquid Methane Test Results of the RS-18 Lunar Ascent Engine at Simulated Altitude Conditions at NASA White Sands Test Facility

Tests were conducted with the RS-18 rocket engine using liquid oxygen (LO2) and liquid methane (LCH4) propellants under simulated altitude conditions at NASA Johnson Space Center White Sands Test Facility (WSTF). This project is part of NASA's Propulsion and Cryogenics Advanced Development (PCAD) project. "Green" propellants, such as LO2/LCH4, offer savings in both performance and safety over equivalently sized hypergolic propulsion systems in spacecraft applications such as ascent engines or service module engines. Altitude simulation was achieved using the WSTF Large Altitude Simulation System, which provided altitude conditions equivalent up to ~122,000 ft (~37 km). For specific impulse calculations, engine thrust and propellant mass flow rates were measured. LO2 flow ranged from 5.9 - 9.5 lbm/sec (2.7 - 4.3 kg/sec), and LCH4 flow varied from 3.0 - 4.4 lbm/sec (1.4 - 2.0 kg/sec) during the RS-18 hot-fire test series. Propellant flow rate was measured using a coriolis mass-flow meter and compared with a serial turbine-style flow meter. Results showed a significant performance measurement difference during ignition startup due to two-phase flow effects. Subsequent cold-flow testing demonstrated that the propellant manifolds must be adequately flushed in order for the coriolis flow meters to give accurate data. The coriolis flow meters were later shown to provide accurate steady-state data, but the turbine flow meter data should be used in transient phases of operation. Thrust was measured using three load cells in parallel, which also provides the capability to calculate thrust vector alignment. Ignition was demonstrated using a gaseous oxygen/methane spark torch igniter. Test objectives for the RS-18 project are 1) conduct a shakedown of the test stand for LO2/methane lunar ascent engines, 2) obtain vacuum ignition data for the torch and pyrotechnic igniters, and 3) obtain nozzle kinetics data to anchor two-dimensional kinetics codes. All of these objectives were met with the RS-18 data and additional testing data from subsequent LO2/methane test programs in 2009 which included the first simulated-altitude pyrotechnic ignition demonstration of LO2/methane.

Melcher, John C., IV↗
NASA NTRS2012

Analysis of Buzz in a Supersonic Inlet

A dual-stream, low-boom supersonic inlet designed for use on a small, Mach 1.6 aircraft was tested experimentally in the 8- by 6-Foot Supersonic Wind Tunnel (SWT) at the NASA Glenn Research Center (GRC). The tests showed that the inlet had good recovery and stable operation over large mass flow range. The inlet went into buzz at mass flows well below that needed for engine operation, and the experiments generated a wealth of data during buzz. High frequency response pressure measurements and high-speed schlieren videos were recorded for many buzz events. The objective of the present work was to use computational fluid dynamics (CFD) to predict some of the experimental data taken during buzz, compare those predictions to the experimental data, and to use both datasets to explain the physics of the buzz cycle. The calculations were done with the Wind-US CFD code using a second-order time-accurate differencing scheme and the SST turbulence model. Computed Mach number contours were compared with schlieren images, and ensemble-averaged unsteady pressures were compared to data. The results showed that the buzz cycle consisted partly of spike buzz, an unsteady oscillation of the main shock at the spike tip while the inlet pressure dropped, and partly of choked flow while the inlet repressurized. Most of the results could be explained by theory proposed by Dailey in 1954, but did not support commonly used acoustic resonance explanations.

Chima, Rodrick V.↗
NASA NTRS2020

Wind Tunnel Testing of High Efficiency Low Power (HELP) Actuation for Active Flow Control

A High Efficiency Low Power (HELP) methodology has been successfully developed and tested on an active flow control (AFC) version of the 10%-scale high-lift Common Research Model (CRM-HL) at the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel (14x22). The AFC variant of the CRM-HL, designated as CRM-SHL-AFC, was integrated with modular HELP actuator cartridges on the shoulder of its highly deflected (≥50°) simplehinged flaps. A reference conventional CRM-HL configuration equipped with Fowler flaps and a nacelle chine was chosen to provide the targeted lift enhancement goals and for comparison with the results of the CRM-SHL-AFC. The current data are presented with the Transonic Wall Interference Correction System (TWICS) method applied. The HELP actuators, which use a combination of unsteady sweeping jets and steady discrete jets in tandem, were designed to overcome strong adverse pressure gradients, while minimizing the pneumatic power usage. The power coefficient (C(sub π)), which takes account of both supply air pressure and mass flow usage for the AFC actuators, is a useful parameter for judging the actuators’ performance efficiency and perhaps for scaling-up of the AFC system. Full HELP actuation coverage with a constant spanwise nozzle pressure ratio (NPR) was able to achieve the lift enhancement goals for the entire lift curve and was found to be the most effective AFC case. For the most effective case, mass flow rates greater than ~0.91 lbm/s and NPR values greater than ~1.8 (corresponding to C(sub π) = 0.18) are needed to achieve the lift enhancement goal at α = 9°, whereas mass flow rates greater than ~1.23 lbm/s and NPR values greater than ~2.3 (corresponding to C(sub π) = 0.3) are needed to achieve the lift enhancement goal at α = 17°. Surface pressure distributions indicate that the HELP actuation on the flap shoulder increased the suction pressures (and flow circulation) globally in both the streamwise and the spanwise directions, and thereby enhanced the lift over the entire high-lift system. The CRM-SHLAFC configuration equipped with HELP actuation was able to match or exceed the lift performance of the reference conventional CRM-HL, thus meeting the objective of the research.

John C Lin↗
NASA NTRS1993

A parametric study of bleed in shock boundary layer interactions

A numerical investigation was conducted to study the effect of bleed configuration on oblique-shock wave/turbulent boundary-layer interactions. Bleed is applied through a normal slot across the shock impingement location. The numerical solution to the compressible Navier-Stokes equations is obtained for the turbulent flow throughout the interaction zone and inside the bleed slot for bleed mass flow rates up to 57 percent of the boundary layer mass flow upstream of the interaction. The results indicate that the bleed slot performance improves as the slot width decreases and the length to width ratio increases. This is reflected as an increase in the bleed discharge coefficient and total pressure, and a reduction in the boundary layer momentum and displacement thickness downstream.

Hamed, A.↗