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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 55 records · Page 3
All sourcesNASA NTRSDOE OSTI
NASA NTRS2003

High-Pressure Transport Properties Of Fluids: Theory And Data From Levitated Drops At Combustion-Relevant Temperatures

Although the high pressure multicomponent fluid conservation equations have already been derived and approximately validated for binary mixtures by this PI, the validation of the multicomponent theory is hampered by the lack of existing mixing rules for property calculations. Classical gas dynamics theory can provide property mixing-rules at low pressures exclusively. While thermal conductivity and viscosity high-pressure mixing rules have been documented in the literature, there is no such equivalent for the diffusion coefficients and the thermal diffusion factors. The primary goal of this investigation is to extend the low pressure mixing rule theory to high pressures and validate the new theory with experimental data from levitated single drops. The two properties that will be addressed are the diffusion coefficients and the thermal diffusion factors. To validate/determine the property calculations, ground-based experiments from levitated drops are being conducted.

Bellan, Josette↗
NASA NTRS1971

Shuttle APS propellant thermal conditioner study

The conditioner design concept selected for evaluation consists of an integral reactor and baffle-type heat exchanger. Heat exchange is accomplished by flowing reactor hot gases past a series of slotted and formed plates, through which the conditioned propellant flows. Heat transfer analysis has resulted in the selection of a reactor hot gas nominal mixture ratio of 1.0, giving a combustion temperature of 1560 F with a hydrogen inlet temperature of 275 R. Worst case conditions result in a combustion gas temperature of 2060 F, satisfying the condition of no damage to the conditioner in case of failure to flow cold fluid. In addition, evaluation of hot gas flow requirements and conditioner weight has resulted in the selection of a reactor hot gas exhaust temperature of 750 R.

Fulton, D. L.↗
NASA NTRS1978

Combuster

A combuster is provided for utilizing a combustible mixture containing fuel and air, to heat a load fluid such as water or air, in a manner that minimizes the formation of nitrogen oxide. The combustible mixture passes through a small diameter tube where the mixture is heated to its combustion temperature, while the load fluid flows past the outside of the tube to receive heat. The tube is of a diameter small enough that the combustible mixture cannot form a flame, and yet is not subject to wall quench, so that combustion occurs, but at a temperature less than under free flame conditions. Most of the heat required for heating the combustible mixture to its combustion temperature, is obtained from heat flow through the walls of the pipe to the mixture.

Mckay, R. A.↗
NASA NTRS1984

Fuel rich catalytic comustion: The first stage of a two-stage combustor

An experimental program demonstrated that fuel-rich catalytic combustion can be accomplished soot free as long as the combustion temperature is less than the temperature at the rich limit of combustion. Although soot was not measured directly, three pieces of data strongly suggest that it was not present: (1) the product gases were completely transparent and produced no radiation characteristic of soot, (2) measured reaction temperatures followed closely those calculated for equilibrium with no soot present, and (3) over 99 percent of the carbon was accounted for in the measured reaction products. Data for two catalyst configurations were taken along with gas samples at two locations downstream of the catalyst bed.

Brabbs, T. A.↗
NASA NTRS1950

Altitude-chamber Performance of British Roll-royce Nene II Engine IV : Effect of Operational Variables on Temperature Distribution at Combustion-chamber Outlets

Temperature surveys were made at the combustion-chamber outlets of a British Rolls-Royce Nene II engine. The highest mean nozzle-vane and mean gas temperatures were found to occur at a radius approximately 75% of the nozzle-vane length from the inner ring of the nozzle-vane assembly. Variations in engine speed, jet-nozzle area, simulated altitude, and simulated flight speed altered the temperature level but did not materially affect the pattern of radial temperature distribution.

COMBUSTION - TURBINE ENGINES↗
NASA NTRS2015

Low-Cost, High-Performance Combustion Chamber

Ultramet designed and fabricated a lightweight, high-temperature combustion chamber for use with cryogenic LOX/CH4 propellants that can deliver a specific impulse of approx.355 seconds. This increase over the current 320-second baseline of nitrogen tetroxide/monomethylhydrazine (NTO/MMH) will result in a propellant mass decrease of 55 lb for a typical lunar mission. The material system was based on Ultramet's proven oxide-iridium/rhenium architecture, which has been hot-fire tested with stoichiometric oxygen/hydrogen for hours. Instead of rhenium, however, the structural material was a niobium or tantalum alloy that has excellent yield strength at both ambient and elevated temperatures. Phase I demonstrated alloys with yield strength-to-weight ratios more than three times that of rhenium, which will significantly reduce chamber weight. The starting materials were also two orders of magnitude less expensive than rhenium and were less expensive than the C103 niobium alloy commonly used in low-performance engines. Phase II focused on the design, fabrication, and hot-fire testing of a 12-lbf thrust class chamber with LOX/CH4, and a 100-lbf chamber for LOX/CH4. A 5-lbf chamber for NTO/MMH also was designed and fabricated.

Fortini, Arthur J.↗
NASA NTRS1980

Development of a solvent processed insensitive propellant

Two types of low vulnerability propellants are studied which are distinguished by whether the binder is a rubber, such as polyurethane or CTBN, or a plasticizable polymer such as ethyl cellulose or cellulose acetate. The former propellants are made by a partial cure extrusion process while the latter are made by the conventional solvent process. Emphasis is given to a cellulose binder (plasticizer) RDX composition. The type of binder used, the particle size of the RDX and the presence of small quantities of nitrocellulose in the solvent processed compositions have important influences on the mechanical and combustion characteristics of the propellant. The low temperature combustion is of particular concern because of potential breakup of the grains that can lead to instability.

Trask, R.↗
NASA NTRS1993

Combustion synthesis of advanced composite materials

Self-propagating high temperature (combustion) synthesis (SHS), has been investigated as a means of producing both dense and expanded (foamed) ceramic and ceramic-metal composites, ceramic powders and whiskers. Several model exothermic combustion synthesis reactions were used to establish the importance of certain reaction parameters, e.g., stoichiometry, green density, combustion mode, particle size, etc. on the control of the synthesis reaction, product morphology and properties. The use of an in situ liquid infiltration technique and the effect of varying the reactants and their stoichiometry to provide a range of reactant and product species i.e., solids, liquids and gases, with varying physical properties e.g., volatility and thermal conductivity, on the microstructure and morphology of synthesized composite materials is discussed. Conducting the combustion synthesis reaction in a reactive gas environment to take advantage of the synergistic effects of combustion synthesis and vapor phase transport is also examined.

Moore, John J.↗
DOE OSTI2026

Flame dynamics and kinetic coupling of ammonia and dimethyl-ether in non-premixed cool and warm flames at elevated pressure

Developing advanced low-temperature combustion engines with ammonia-biofuel blends requires a comprehensive understanding of low-temperature flame dynamics and kinetic interactions between ammonia and oxygenated fuels at elevated pressures. This work aims to study the dynamics and kinetics of non-premixed Dimethyl Ether (DME)/Ammonia (NH 3 ) cool and warm flames, and their reignition to hot flames. A counterflow burner is employed to establish DME/NH 3 cool/warm flames at pressures up to 5 atm. The extinction limits of cool flame and the reignition limits of warm flame to hot flame are measured by varying NH 3 concentrations and compared to simulations to quantitatively examine the effects on DME/NH 3 flames. It is found that NH 3 inhibits low-temperature DME oxidation and results in lower cool flame extinction limits. Warm flames in the presence of NH 3 are observed for the first time, revealing a non-monotonic effect of NH 3 addition: a small amount of NH 3 presence enhances warm flame chemistry and promotes reignition to hot flames, while a high NH 3 concentration weakens the warm flame. This trend is further explained by 0-D PSR kinetic simulations and 1-D S-curve flame dynamic calculations. Three flame transition regimes between cool flames (CF), warm flames (WF), and hot flames (HF) by different levels of NH 3 additions at a specific strain rate are identified, namely WFHF reignition, WF-CF transition, and WF extinction. Reaction sensitivity analyses of OH at low temperatures show that NH3 inhibits DME oxidation through OH consumption via H-abstraction and the kinetic couplings of RO 2 /NH 2 , RO 2 /NO x , R/NO x , and O 2 QOOH/NO x further suppress the low-temperature branching. At intermediate-temperatures, NH 2 /NO x /HO 2 coupling promotes warm flames via the pathway NH 2 → H 2 NO → HNO → NO by converting O 2 → HO 2 → OH. At even higher NH₃ concentrations, radical termination reactions of NH 2 + NO/NO 2 and excessive OH consumption via H-abstraction inhibit the flame. The insights into the kinetic coupling between NH 3 and low-temperature chemistry at elevated pressure and its impact on the dynamics of cool-warm-hot flame transitions will contribute to advancing combustion technologies with reduced emissions and improved energy-efficiency.

42 ENGINEERING↗
NASA NTRS1993

Combustor technology for future small gas turbine aircraft

Future engine cycles proposed for advanced small gas turbine engines will increase the severity of the operating conditions of the combustor. These cycles call for increased overall engine pressure ratios which increase combustor inlet pressure and temperature. Further, the temperature rise through the combustor and the corresponding exit temperature also increase. Future combustor technology needs for small gas turbine engines is described. New fuel injectors with large turndown ratios which produce uniform circumferential and radial temperature patterns will be required. Uniform burning will be of greater importance because hot gas temperatures will approach turbine material limits. The higher combustion temperatures and increased radiation at high pressures will put a greater heat load on the combustor liners. At the same time, less cooling air will be available as more of the air will be used for combustion. Thus, improved cooling concepts and/or materials requiring little or no direct cooling will be required. Although presently there are no requirements for emissions levels from small gas turbine engines, regulation is expected in the near future. This will require the development of low emission combustors. In particular, nitrogen oxides will increase substantially if new technologies limiting their formation are not evolved and implemented. For example, staged combustion employing lean, premixed/prevaporized, lean direct injection, or rich burn-quick quench-lean burn concepts could replace conventional single stage combustors.

Lyons, Valerie J.↗
NASA NTRS1994

Combustor technology for future small gas turbine aircraft

To enhance fuel efficiency, future advanced small gas turbine engines will utilize engine cycles calling for overall engine pressure ratios, leading to higher combustor inlet pressures and temperatures. Further, the temperature rise through the combustor and the corresponding exit temperature are also expected to increase. This report describes future combustor technology needs for small gas turbine engines. New fuel injectors with large turndown ratios which produce uniform circumferential and radial temperature patterns will be required. Uniform burning will be of greater importance because hot gas temperatures will approach turbine material limits. The higher combustion temperatures and increased radiation at high pressures will put a greater heat load on the combustor liners. At the same time, less cooling air will be available as more of the air will be used for combustion. Thus, improved cooling concepts and/or materials requiring little or no direct cooling will be required. Although presently there are no requirements for emissions levels from small gas turbine engines, regulation is anticipated in the near future. This will require the development of low emission combustors. In particular, nitrogen oxides will increase substantially if new technologies limiting their formation are not evolved and implemented. For example, staged combustion employing lean, premixed/prevaporized, lean direct injection, or rich burn-quick quench-lean burn concepts could replace conventional single stage combustors. Due to combustor size considerations, staged combustion is more easily accommodated in large engines. The inclusion of staged combustion in small engines will pose greater combustor design challenges.

Lyons, Valerie J.↗
NASA NTRS1948

Photographic Study of Combustion in a Rocket Engine I : Variation in Combustion of Liquid Oxygen and Gasoline with Seven Methods of Propellant Injection

Motion pictures at camera speeds up to 3000 frames per second were taken of the combustion of liquid oxygen and gasoline in a 100-pound-thrust rocket engine. The engine consisted of thin contour and injection plates clamped between two clear plastic sheets forming a two-dimensional engine with a view of the entire combustion chamber and nozzle. A photographic investigation was made of the effect of seven methods of propellant injection on the uniformity of combustion. From the photographs, it was found that the flame front extended almost to the faces of the injectors with most of the injection methods, all the injection systems resulted in a considerable nonuniformity of combustion, and luminosity rapidly decreased in the divergent part of the nozzle. Pressure vibration records indicated combustion vibrations that approximately corresponded to the resonant frequencies of the length and the thickness of the chamber. The combustion temperature divided by the molecular weight of the combustion gases as determined from the combustion photographs was about 50 to 70 percent of the theoretical value.

Bellman, Donald R↗
NASA NTRS2001

Combustion Synthesis of Glass-Ceramic Composites Under Terrestrial and Reduced Gravity Conditions

Glasses based on B2O3-Al2O3-BaO-and B2O3-Al2O3-MgO have been produced by the combustion synthesis technique. The combustion temperature, wave velocity for selected compositions are presented. Combustion reactions of these materials were typically low exothermic, resulting in unstable combustion waves. Microstructural characterization of these materials indicated that the glass formation region was similar to those that were produced by the traditional technique. Results of the effect of gravity on the glass formation (or divitrification) studied onboard of KC-135 is also presented.

Manerbino, Anthony↗
NASA NTRS1998

Feasibility Assessment of Thermal Barrier Seals for Extreme Transient Temperatures

The assembly joints of modem solid rocket motor cases are generally sealed using conventional O-ring type seals. The 5500+ F combustion gases produced by rocket motors are kept a safe distance away from the seals by thick layers of phenolic insulation. Special compounds are used to fill insulation gaps leading up to the seals to prevent a direct flowpath to them. Design criteria require that the seals should not experience torching or charring during operation, or their sealing ability would be compromised. On limited occasions, NASA has observed charring of the primary O-rings of the Space Shuttle solid rocket nozzle assembly joints due to parasitic leakage paths opening up in the gap-fill compounds during rocket operation. NASA is investigating different approaches for preventing torching or charring of the primary O-rings. One approach is to implement a braided rope seal upstream of the primary O-ring to serve as a thermal barrier that prevents the hot gases from impinging on the O-ring seals. This paper presents flow, resiliency, and thermal resistance for several types of NASA rope seals braided out of carbon fibers. Burn tests were performed to determine the time to burn through each of the seals when exposed to the flame of an oxyacetylene torch (5500 F), representative of the 5500 F solid rocket motor combustion temperatures. Rope seals braided out of carbon fibers endured the flame for over six minutes, three times longer than solid rocket motor burn time. Room and high temperature flow tests are presented for the carbon seals for different amounts of linear compression. Room temperature compression tests were performed to assess seal resiliency and unit preloads as a function of compression. The thermal barrier seal was tested in a subscale "char" motor test in which the seal sealed an intentional defect in the gap insulation. Temperature measurements indicated that the seal blocked 2500 F combustion gases on the upstream side with very little temperature rise on the downstream side.

Steinetz, Bruce M.↗
NASA NTRS1981

Performance calculations for 1000 MWe MHD/steam power plants

The effects of MHD generator operating conditions and constraints on the performance of MHD/steam power plants are investigated. Power plants using high temperature combustion air preheat (2500 F) and plants using intermediate temperature preheat (1100 F) with oxygen enrichment are considered. Variations of these two types of power plants are compared on the basis of fixed total electrical output (1000 MWe). Results are presented to show the effects of generator plant length and level of oxygen enrichment on the plant thermodynamic efficiency and on the required generator mass flow rate. Factors affecting the optimum levels of oxygen enrichment are analyzed. It is shown that oxygen enrichment can reduce magnet stored energy requirement.

Pian, C. C. P.↗
NASA NTRS1994

Making Ceramic Components For Advanced Aircraft Engines

Lightweight, oxidation-resistant silicon nitride components containing intricate internal cooling and hydraulic passages and capable of withstanding high operating temperatures made by ceramic-platelet technology. Used to fabricate silicon nitride test articles of two types: components of methane-cooled regenerator for air turbo ramjet engine and components of bipropellant injector for rocket engine. Procedures for development of more complex and intricate components established. Technology has commercial utility in automotive, aircraft, and environmental industries for manufacture of high-temperature components for use in regeneration of fuels, treatment of emissions, high-temperature combustion devices, and application in which other high-temperature and/or lightweight components needed. Potential use in fabrication of combustors and high-temperature acoustic panels for suppression of noise in future high-speed aircraft.

Franklin, J. E.↗
NASA NTRS2006

Combustion Synthesis of Ca3(PO4)2 Net-Shape Surgical Implants

Self-propagating high-temperature combustion synthesis (SHS) is the basis of a method of making components of porous tricalcium phosphate [Ca3(PO4)2] and related compounds in net sizes and shapes for use as surgical implants that are compatible with bone. The SHS method offers advantages over prior methods of manufacturing Ca3(PO4)2-based surgical implants.

Ayers, Reed A.↗