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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 289 records · Page 16

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NASA NTRS1994

Solid rocket combustion simulator technology using the hybrid rocket for simulation

The hybrid rocket is reexamined in light of several important unanswered questions regarding its performance. The well-known heat transfer limited burning rate equation is quoted, and its limitations are pointed out. Several inconsistencies in the burning rate determination through fuel depolymerization are explicitly discussed. The resolution appears to be through the postulate of (surface) oxidative degradation of the fuel. Experiments are initiated to study the fuel degradation in mixtures of nitrogen/oxygen in the 99.9 percent/0.1 percent to 98 percent/2 percent range. The overall hybrid combustion behavior is studied in a 2 in-diameter rocket motor, where a PMMA tube is used as the fuel. The results include detailed, real-time infrared video images of the combustion zone. Space- and time-averaged images give a broad indication of the temperature reached in the gases. A brief outline is shown of future work, which will specifically concentrate on the exploration of the role of the oxidizer transport to the fuel surface, and the role of the unburned fuel that is reported to escape below the classical time-averaged boundary layer flame.

Ramohalli, Kumar↗

NASA NTRS1994

WINCLR: a Computer Code for Heat Transfer and Clearance Calculation in a Compressor

One of the concerns during inclement weather operation of aircraft in rain and hail storm conditions is the nature and extent of changes in compressor casing clearance. An increase in clearance affects efficiency while a decrease may cause blade rubbing with the casing. The change in clearance is the result of geometrical dimensional changes in the blades, the casing and the rotor due to heat transfer between those parts and the two-phase working fluid. The heat transfer interacts nonlinearly with the performance of the compressor, and, therefore, the determination of clearance changes necessitates a simultaneous determination of change in performance of the compressor. A computer code the WINCLR has been designed for the determination of casing clearance, that is operated interactively with the PURDU-WINCOF I code designed previously for determining the performance of a compressor. A detailed description of the WINCLR code is provided in a companion report. The current report provides details of the code with an illustrative example of application to the case of a multistage compressor. It is found in the example case that under given ingestion and operational conditions, it is possible for a compressor to undergo changes in performance in the front stages and rubbing in the back stages.

Bose, T. K.↗

NASA NTRS1996

Predicted Turbine Heat Transfer for a Range of Test Conditions

Comparisons are shown between predictions and experimental data for blade and endwall heat transfer. The comparisons of computational domain parisons are given for both vane and rotor geometries over an extensive range of Reynolds and Mach numbers. Comparisons are made with experimental data from a variety of sources. A number of turbulence models are available for predicting blade surface heat transfer, as well as aerodynamic performance. The results of an investigation to determine the turbulence model which gives the best agreement with experimental data over a wide range of test conditions are presented.

Boyle, R. J.↗

NASA NTRS2014

Co-Optimization of Blunt Body Shapes for Moving Vehicles

A method and associated system for multi-disciplinary optimization of various parameters associated with a space vehicle that experiences aerocapture and atmospheric entry in a specified atmosphere. In one embodiment, simultaneous maximization of a ratio of landed payload to vehicle atmospheric entry mass, maximization of fluid flow distance before flow separation from vehicle, and minimization of heat transfer to the vehicle are performed with respect to vehicle surface geometric parameters, and aerostructure and aerothermal vehicle response for the vehicle moving along a specified trajectory. A Pareto Optimal set of superior performance parameters is identified.

Brown, James L.↗

NASA NTRS2016

Aerogel Hybrid Composite Materials: Designs and Testing for Multifunctional Applications

This webinar will introduce the broad spectrum of aerogel composites and their diverse performance properties such as reduced heat transfer to energy storage, and expands specifically on the aerogel/fiber laminate systems and testing methodologies. The multi-functional laminate composite system, AeroFiber, and its construction is designed by varying the type of fiber (e.g. polyester, carbon, Kevlar®, Spectra® or Innegral(TradeMark) and combinations thereof), the aerogel panel type and thickness, and overall layup configuration. The combination and design of materials may be customized and tailored to achieve a range of desired properties in the resulting laminate system. Multi-functional properties include structural strength, impact resistance, reduction in heat transfer, increased fire resistance, mechanical energy absorption, and acoustic energy dampening. Applications include aerospace, aircraft, automotive, boating, building and construction, lightweight portable structures, liquefied natural gas, cryogenics, transportation and energy, sporting equipment, and military protective gear industries.

Hybrid↗

NASA NTRS2015

The Effect of Core Configuration on Thermal Barrier Thermal Performance

Thermal barriers and seals are integral components in the thermal protection systems (TPS) of nearly all aerospace vehicles. They are used to minimize heat transfer through interfaces and gaps and protect underlying temperature-sensitive components. The core insulation has a significant impact on both the thermal and mechanical properties of compliant thermal barriers. Proper selection of an appropriate core configuration to mitigate conductive, convective and radiative heat transfer through the thermal barrier is challenging. Additionally, optimization of the thermal barrier for thermal performance may have counteracting effects on mechanical performance. Experimental evaluations have been conducted to better understand the effect of insulation density on permeability and leakage performance, which can significantly impact the resistance to convective heat transfer. The effect of core density on mechanical performance was also previously investigated and will be reviewed. Simple thermal models were also developed to determine the impact of various core parameters on downstream temperatures. An extended understanding of these factors can improve the ability to design and implement these critical TPS components.

flow tests↗

NASA NTRS2000

Effect of Elevated Pressure on the Heat Transfer and Power Requirements During Bridgman Growth of PMN-PT Crystals

Performance of the furnace during Bridgman growth of the lead magnesium niobate-lead titanate crystal (PMN-PT) is analyzed. PMN-PT is electrostrictive ceramic that has near ideal strain-voltage function. Furthermore piezoelectric (2000 to 2300 pC/N) and coupling (92 to 95%) constants are exceptionally good. Due to these properties PMN-PT has wide range of applications - from sonars to transducers in a high precision optical systems. In this research first attempt to crystallize PMN-PT in a Mellen type vertical Bridgman furnace was not successful, as melting temperature of precursor materials was not achieved. At this point choice was between building a new more powerful facility or finding ways to enhance performance of the existing furnace. Besides adjusting power supply to the individual heating elements, redesigning ampoule holding cartridge and improving furnace insulation one more radical improvement was proposed. The entire furnace was placed into the high pressure chamber. Further experiments confirmed that temperature inside the furnace was increased sufficiently to melt precursor materials to obtain PMN-PT. Numerical modeling is undertaken to find limitations of this technique and to predict temperature distribution inside the ampoule. It is of interest also to account for main factors contributing to a higher temperatures achieved in the furnace under the higher pressure (up to 10 atm.). Numerical model of the furnace is based on general purpose finite - element code FIDAP and on previous efforts to model Bridgman type furnace with multiply heaters. In order to account for all heat transfer mechanism involved - conduction, convection and radiation - different parts of the furnace are modeled in accordance with expected dominant mode of heat transfer - conduction in the solid parts, conduction and radiation in the ampoule, gas convection and conduction in the furnace openings complemented with wall-to-wall radiation. Because of these complicating factors, dimensional rather than non-dimensional modeling is performed using steady-state 2-D and 3-D models. Particular attention is paid to the modeling of radiation in a semitransparent material of ampoule 7 sapphire. The radiation model is validated by solving realistic test problem - conduction and radiation heat transfer in the fused quartz. Results are in agreement with both experimental and analytical data.

Bune, Andris↗

NASA NTRS2006

Analysis of a Cylindrical Specimen Heated by an Impinging Hot Hydrogen Jet

A computational conjugate heat transfer methodology was developed, as a first step towards an efficient and accurate multiphysics, thermo-fluid computational methodology to predict environments for hypothetical solid-core, nuclear thermal engine thrust chamber and components. A solid conduction heat transfer procedure was implemented onto a pressure-based, multidimensional, finite-volume, turbulent, chemically reacting, thermally radiating, and unstructured grid computational fluid dynamics formulation. The conjugate heat transfer of a cylindrical material specimen heated by an impinging hot hydrogen jet inside an enclosed test fixture was simulated and analyzed. The solid conduction heat transfer procedure was anchored with a standard solid heat transfer code. Transient analyses were then performed with ,variable thermal conductivities representing three composites of a material utilized as flow element in a legacy engine test. It was found that material thermal conductivity strongly influences the transient heat conduction characteristics. In addition, it was observed that high thermal gradient occur inside the cylindrical specimen during an impulsive or a 10 s ramp start sequence, but not during steady-state operations.

Wang, Ten-See↗

NASA NTRS1982

Heat-Energy Analysis for Solar Receivers

Heat-energy analysis program (HEAP) solves general heat-transfer problems, with some specific features that are "custom made" for analyzing solar receivers. Can be utilized not only to predict receiver performance under varying solar flux, ambient temperature and local heat-transfer rates but also to detect locations of hotspots and metallurgical difficulties and to predict performance sensitivity of neighboring component parameters.

Lansing, F. L.↗

NASA NTRS1977

Evaluation of thermal soak times

A mathematical model for the heat transfer within the electronics package of a Chaparral missile was performed. The Grashof number for this configuration was less than 2000 which indicated that the primary mode of heat transfer was conduction. The Vodicka theory for heat conduction in laminated composite media was utilized to obtain the solution for the model.

Mulholland, G. P.↗

NASA NTRS1971

Design and preliminary testing of a Brayton space radiator concept

A low-temperature radiator has been designed to reject 17.62 kW of thermal power from an isotope Brayton power system. This power system has an electrical output of 7.1 kW. The total radiator consists of two radiators connected in series. Both radiators have a redundant coolant loop, shared fins, and are close to flight configuration. A silicone oil is used as the working fluid. Low flow rates and moderate pressure drop in the radiator required a small hydraulic diameter with large heat transfer area and large flow area. To obtain these characteristics and meet schedule limitations, an expanded tube configuration (with an approximately trapezoidal cross section) was selected. Because of the low flow rates and unconventional tube shape, small scale tests were performed to evaluate the validity of existing heat transfer and pressure drop correlations to this case. The results of these tests were in good agreement with predictions.

Miller, T. J.↗

NASA NTRS1991

Heat pipe radiation cooling of advanced hypersonic propulsion system components

Heat transfer, heat pipe, and system studies were performed to assess the newly proposed heat pipe radiation cooling (HPRC) concept. With an HPRC system, heat is removed from the ramburner and nozzle of a hypersonic aircraft engine by a surrounding, high-temperature, heat pipe nacelle structure, transported to nearby external surfaces, and rejected to the environment by thermal radiation. With HPRC, the Mach number range available for using hydrocarbon fuels for aircraft operation extends into the Mach 4 to Mach 6 range, up from the current limit of about Mach 4. Heat transfer studies using a newly developed HPRC computer code determine cooling system and ramburner and nozzle temperatures, heat loads, and weights for a representative combined-cycle engine cruising at Mach 5 at 80,000 ft altitude. Heat pipe heat transport calculations, using the Los Alamos code HTPIPE, reveal that adequate heat trasport capability is available using molybdenum-lithium heat pipe technology. Results show that the HPRC system radiator area is limited in size to the ramburner-nozzle region of the engine nacelle; reasonable system weights are expected; hot section temperatures are consistent with advanced structural materials development goals; and system impact on engine performance is minimal.

Martin, R. A.↗

NASA NTRS2012

NARloy-Z-Carbon Nanotube Composites

Motivation: (1) NARloy-Z (Cu-3%Ag-0.5%Zr) is the state of the art, high thermal conductivity structural alloy used for making liquid rocket engine main combustion chamber liner. It has a Thermal conductivity approx 80% of pure copper. (2) Improving the thermal conductivity of NARloy-Z will help to improve the heat transfer efficiency of combustion chamber. (3)Will also help to reduce the propulsion system mass and increase performance. It will also increases thrust to weight ratio. (4) Improving heat transfer helps to design and build better thermal management systems for nuclear propulsion and other applications. Can Carbon nanotubes (CNT) help to improve the thermal conductivity (TC)of NARloy-Z? (1)CNT's have TC of approx 20X that of copper (2) 5vol% CNT could potentially double the TC of NARloy-Z if properly aligned (3) Improvement will be less if CNT s are randomly distributed, provided there is a good thermal bond between CNT and matrix. Prior research has shown poor results (1) No TC improvement in the copper-CNT composite reported (2)Reported values are typically lower (3) Attributed to high contact thermal resistance between CNT and Cu matrix (4)Results suggest that a bonding material between CNT and copper matrix is required to lower the contact thermal resistance It is hypothesized that Zr in NARloy-Z could act as a bonding agent to lower the contact thermal resistance between CNT and matrix.

Bhat, Biliyar N.↗

NASA NTRS1987

Performance and combustion characteristics of direct-injection stratified-charge rotary engines

Computer simulations of the direct-injection stratified-charge (DISC) Wankel engine have been used to calculate heat release rates and performance and efficiency characteristics of the 1007R engine. Engine pressure data have been used in a heat release analysis to study the effects of heat transfer, leakage, and crevice flows. Predicted engine performance data are compared with experimental test data over a range of engine speeds and loads. An examination of methods to improve the performance of the Wankel engine with faster combustion, reduced leakage, higher compression ratio, and turbocharging is presented.

Nguyen, Hung Lee↗

NASA NTRS1992

Ignition and combustion characteristics of metallized propellants

During this reporting period, theoretical work on the secondary atomization process was continued and the experimental apparatus was improved. A one-dimensional model of a rocket combustor, incorporating multiple droplet size classes, slurry combustion, secondary atomization, radiation heat transfer, and two-phase slip between slurry droplets and the gas flow was derived and a computer code was written to implement this model. The STANJAN chemical equilibrium solver was coupled with this code to yield gas temperature, density, and composition as functions of axial location. Preliminary results indicate that the model is performing correctly, given current model assumptions. Radiation heat transfer in the combustion chamber is treated as an optically-thick participating media problem requiring a solution of the radiative transfer equation. A cylindrical P sub 1 approximation was employed to yield an analytical expression for chamber-wall heat flux at each axial location. The code exercised to determine the effects of secondary atomization intensity, defined as the number of secondary drops produced per initial drop, on chamber burnout distance and final Al2O3 agglomerate diameter. These results indicate that only weak secondary atomization is required to significantly reduce these two parameters. Stronger atomization intensities were found to yield decreasing marginal benefits. The experimental apparatus was improved to reduce building vibration effects on the optical system alignment. This was accomplished by mounting the burner and the transmitting/receiving optics on a single frame supported by vibration-isolation legs. Calibration and shakedown tests indicate that vibration problems were eliminated and that the system is performing correctly.

Mueller, D. C.↗

NASA NTRS1976

Heat transfer evaluation in a plasma core reactor

Numerical evaluations of heat transfer in a fissioning uranium plasma core reactor cavity, operating with seeded hydrogen propellant, was performed. A two-dimensional analysis is based on an assumed flow pattern and cavity wall heat exchange rate. Various iterative schemes were required by the nature of the radiative field and by the solid seed vaporization. Approximate formulations of the radiative heat flux are generally used, due to the complexity of the solution of a rigorously formulated problem. The present work analyzes the sensitivity of the results with respect to approximations of the radiative field, geometry, seed vaporization coefficients and flow pattern. The results present temperature, heat flux, density and optical depth distributions in the reactor cavity, acceptable simplifying assumptions, and iterative schemes. The present calculations, performed in cartesian and spherical coordinates, are applicable to any most general heat transfer problem.

Smith, D. E.↗

NASA NTRS1953

An Investigation of High-frequency Combustion Oscillations in Liquid-propellant Rocket Engines

An experimental investigation of high-frequency combustion oscillations (screaming) was conducted with a 100-pound-thrust acid-hydrocarbon rocket engine and a 500-pound-thrust oxygen-fuel rocket engine. The oscillation frequencies could be correlated as a linear function of the parameter C/L, where C is the experimentally measured characteristic velocity and L is the combustion-chamber length. The tendency of the engines to scream increased as chamber length was increased. With engine configurations that normally had a low efficiency, screaming resulted in increased performance; at the same time, a five to tenfold increase in heat-transfer rate occurred. It was possible, however, to achieve good performance without screaming.

COMBUSTION - EFFECT OF ENGINE OPERATING CONDITIONS↗

NASA NTRS1989

Transient hot-film sensor response in a shock tube

Shock tube experiments were performed to determine the response of a hot-film sensor, mounted flush on the side wall of a shock tube, to unsteady flow behind a normal shock wave. The present experiments attempt to isolate the response of the anemometer due only to the change in convective heat transfer at the hot-film surface. The experiments, performed at low supersonic shock speeds in air, are described along with the data acquisition procedure. The change in convective heat transfer is deduced from the data and the results are compared with those from transient boundary layer theory and another set of experimental results. Finally, a transient local heat transfer coefficient is formulated for use as the forcing function in a hot-film sensor instrument model simulation.

Roberts, A. S., Jr.↗