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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 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS2005

Heat Transfer and Thermal Stability Research for Advanced Hydrocarbon Fuel Technologies

In recent years there has been increased interest in the development of a new generation of high performance boost rocket engines. These efforts, which will represent a substantial advancement in boost engine technology over that developed for the Space Shuttle Main Engines in the early 1970s, are being pursued both at NASA and the United States Air Force. NASA, under its Space Launch Initiative s Next Generation Launch Technology Program, is investigating the feasibility of developing a highly reliable, long-life, liquid oxygen/kerosene (RP-1) rocket engine for launch vehicles. One of the top technical risks to any engine program employing hydrocarbon fuels is the potential for fuel thermal stability and material compatibility problems to occur under the high-pressure, high-temperature conditions required for regenerative fuel cooling of the engine combustion chamber and nozzle. Decreased heat transfer due to carbon deposits forming on wetted fuel components, corrosion of materials common in engine construction (copper based alloys), and corrosion induced pressure drop increases have all been observed in laboratory tests simulating rocket engine cooling channels. To mitigate these risks, the knowledge of how these fuels behave in high temperature environments must be obtained. Currently, due to the complexity of the physical and chemical process occurring, the only way to accomplish this is empirically. Heated tube testing is a well-established method of experimentally determining the thermal stability and heat transfer characteristics of hydrocarbon fuels. The popularity of this method stems from the low cost incurred in testing when compared to hot fire engine tests, the ability to have greater control over experimental conditions, and the accessibility of the test section, facilitating easy instrumentation. These benefits make heated tube testing the best alternative to hot fire engine testing for thermal stability and heat transfer research. This investigation used the Heated Tube Facility at the NASA Glenn Research Center to perform a thermal stability and heat transfer characterization of RP-1 in an environment simulating that of a high chamber pressure, regenerative cooled rocket engine. The first step in the research was to investigate the carbon deposition process of previous heated tube experiments by performing scanning electron microscopic analysis in conjunction with energy dispersive spectroscopy on the tube sections. This analysis gave insight into the carbon deposition process and the effect that test conditions played in the formation of deleterious coke. Furthermore, several different formations were observed and noted. One other crucial finding of this investigation was that in sulfur containing hydrocarbon fuels, the interaction of the sulfur components with copper based wall materials presented a significant corrosion problem. This problem in many cases was more life limiting than those posed by the carbon deposition process. The results of this microscopic analysis was detailed and presented at the December 2003 JANNAF Air-Breathing Propulsion Meeting as a Materials Compatibility and Thermal Stability Analysis of common Hydrocarbon Fuels (reference 1).

DeWitt, Kenneth↗
NASA NTRS1979

Some aspects of aircraft jet engine fuels

Technologies are reviewed for improving the thermal stability of jet fuels, with reference to the overheating of fuel tanks in supersonic aircraft. Consideration is given to the development of a jet fuel with high thermal stability by the Polish petroleum industry.

Bekiesinski, R.↗
NASA NTRS2017

Characterization and Detailed Analysis of Regression Behavior for HTPB Solid Fuels Containing High Aluminum Loadings

NASA Marshall Space Flight Center's Materials and Processes Department, with support from the Propulsion Systems Department, has renewed the development and maintenance of a hybrid test bed for exposing ablative thermal protection materials to an environment similar to that seen in solid rocket motors (SRM). The Solid Fuel Torch (SFT), operated during the Space Shuttle program, utilized gaseous oxygen for oxidizer and an aluminized hydroxyl-terminated polybutadiene (HTPB) fuel grain to expose a converging section of phenolic material to a 400 psi, 2-phase flow combustion environment. The configuration allows for up to a 2 foot long, 5 inch diameter fuel grain cartridge. Wanting to now test rubber insulation materials with a turn-back feature to mimic the geometry of an aft dome being impinged by alumina particles, the throat area has now been increased by several times to afford flow similarity. Combined with the desire to maintain a higher operating pressure, the oxidizer flow rate is being increased by a factor of 10. Out of these changes has arisen the need to characterize the fuel/oxidizer combination in a higher mass flux condition than has been previously tested at MSFC, and at which the literature has little to no reporting as well. For (especially) metalized fuels, hybrid references have pointed out possible dependence of fuel regression rate on a number of variables: mass flux, G - oxidizer only (G0), or - total mass flux (Gtot), Length, L, Pressure, P, and Diameter, D.

Kibbey, Timothy P.↗
NASA NTRS1992

Mixing in the dome region of a staged gas turbine combustor

To lower NO(x) emissions from gas-turbine engines the effect of dome design and operational changes on the mixing quality in the fuel-rich region is studied. A statistical analysis is employed to establish the parametric sensitivity in this complex flow. A mixing-effectiveness index is defined and used to optimize the gas-species uniformity and the extent of reaction at the exit plane of the dome. Mixing effectiveness is tied to the fuel and air injection locations, the macroscale structure of the dome aerodynamics, and the level of turbulence. Increases in nozzle/air to fuel ratio, reference velocities, and the dome expansion angle increased the level of turbulence. The optimum configuration featured counter-swirling fuel and air streams and produced a strong torroidal recirculation zone, an effective spray angle of 45 degrees, and azimuthal velocities that decayed to zero inside of two duct diameters. The results underscore the system specific nature of mixing optimization.

Sowa, W. A.↗
NASA NTRS1979

Near-term hybrid vehicle program, phase 1. Appendix D: Sensitivity analysis resport

Parametric analyses, using a hybrid vehicle synthesis and economics program (HYVELD) are described investigating the sensitivity of hybrid vehicle cost, fuel usage, utility, and marketability to changes in travel statistics, energy costs, vehicle lifetime and maintenance, owner use patterns, internal combustion engine (ICE) reference vehicle fuel economy, and drive-line component costs and type. The lowest initial cost of the hybrid vehicle would be $1200 to $1500 higher than that of the conventional vehicle. For nominal energy costs ($1.00/gal for gasoline and 4.2 cents/kWh for electricity), the ownership cost of the hybrid vehicle is projected to be 0.5 to 1.0 cents/mi less than the conventional ICE vehicle. To attain this ownership cost differential, the lifetime of the hybrid vehicle must be extended to 12 years and its maintenance cost reduced by 25 percent compared with the conventional vehicle. The ownership cost advantage of the hybrid vehicle increases rapidly as the price of fuel increases from $1 to $2/gal.

Source record↗
NASA NTRS1995

Dual-Fuel Propulsion in Single-Stage Advanced Manned Launch System Vehicle

As part of the United States Advanced Manned Launch System study to determine a follow-on, or complement, to the Space Shuttle, a reusable single-stage-to-orbit concept utilizing dual-fuel rocket propulsion has been examined. Several dual-fuel propulsion concepts were investigated. These include: a separate-engine concept combining Russian RD-170 kerosene-fueled engines with space shuttle main engine-derivative engines: the kerosene- and hydrogen-fueled Russian RD-701 engine; and a dual-fuel, dual-expander engine. Analysis to determine vehicle weight and size characteristics was performed using conceptual-level design techniques. A response-surface methodology for multidisciplinary design was utilized to optimize the dual-fuel vehicles with respect to several important propulsion-system and vehicle design parameters, in order to achieve minimum empty weight. The tools and methods employed in the analysis process are also summarized. In comparison with a reference hydrogen- fueled single-stage vehicle, results showed that the dual-fuel vehicles were from 10 to 30% lower in empty weight for the same payload capability, with the dual-expander engine types showing the greatest potential.

Lepsch, Roger A., Jr.↗
NASA NTRS2018

Fuel Sensitivity of Gas Emissions, Lean Blowout and Combustion Dynamics for a 9-Point LDI Combustor

Fuel sensitivity of gaseous emissions, approach to lean blowout and combustion dynamics are evaluated in this study. Experiments were conducted at the NASA Glenn Research Center's CE-5 flame tube test facility with a 9-point Swirl-Venturi Lean Direct Injection (SV-LDI) combustor. A reference jet fuel (A2) and two test fuels (C1 and C3) from were provided by the National Jet Fuels Combustion Program (NJFCP). C1 is essentially a 2-component iso-paraffin test fuel with a low cetane number of 17, and C3 is a high viscosity test fuel. Approach to lean blowout was monitored in terms of the rapid increase in CO emissions index as equivalence ratio decreased, but testing did not proceed all the way to lean blowout (LBO). Burning C1 was found to produce lower NOx emissions, but C1 flame temperatures were about 25 K higher relative to A2 at near LBO points (where CO emissions increased very rapidly). The NOx emissions of C3 were similar to A2. At low power conditions where fuel injector performance is not optimized for this 9-point LDI combustor, C3 had higher CO emissions than A2 and C1, likely due to C3's higher viscosity relative to A2 and C1. No discernable difference in combustion dynamics was observed between the three fuels tested in the 9-point LDI combustor. While a systematic ignition test campaign was not conducted, it was observed that C1 required a higher equivalence ratio and inlet air temperature for test rig ignition compared to A2 and C3.

Lean Direct Injection↗
NASA NTRS2018

Fuel Sensitivity of Gas Emissions, Lean Blowout and Combustion Dynamics for a 9-Point LDI Combustor

Fuel sensitivity of gaseous emissions, approach to lean blowout and combustion dynamics are evaluated in this study. Experiments were conducted at the NASA Glenn Research Center's CE-5 flame tube test facility with a 9-point Swirl-Venturi Lean Direct Injection (SV-LDI) combustor. A reference jet fuel (A2) and two test fuels (C1 and C3) from were provided by the National Jet Fuels Combustion Program (NJFCP). C1 is essentially a 2-component iso-paraffin test fuel with a low cetane number of 17, and C3 is a high viscosity test fuel. Approach to lean blowout was monitored in terms of the rapid increase in CO emissions index as equivalence ratio decreased, but testing did not proceed all the way to lean blowout (LBO). Burning C1 was found to produce lower NOx emissions, but C1 flame temperatures were about 25 K higher relative to A2 at near LBO points (where CO emissions increased very rapidly). The NOx emissions of C3 were similar to A2. At low power conditions where fuel injector performance is not optimized for this 9-point LDI combustor, C3 had higher CO emissions than A2 and C1, likely due to C3's higher viscosity relative to A2 and C1. No discernable difference in combustion dynamics was observed between the three fuels tested in the 9-point LDI combustor. While a systematic ignition test campaign was not conducted, it was observed that C1 required a higher equivalence ratio and inlet air temperature for test rig ignition compared to A2 and C3.

Combustion dynamics↗
NASA NTRS1981

Parametric analysis of closed cycle magnetohydrodynamic (MHD) power plants

A parametric analysis of closed cycle MHD power plants was performed which studied the technical feasibility, associated capital cost, and cost of electricity for the direct combustion of coal or coal derived fuel. Three reference plants, differing primarily in the method of coal conversion utilized, were defined. Reference Plant 1 used direct coal fired combustion while Reference Plants 2 and 3 employed on site integrated gasifiers. Reference Plant 2 used a pressurized gasifier while Reference Plant 3 used a ""state of the art' atmospheric gasifier. Thirty plant configurations were considered by using parametric variations from the Reference Plants. Parametric variations include the type of coal (Montana Rosebud or Illinois No. 6), clean up systems (hot or cold gas clean up), on or two stage atmospheric or pressurized direct fired coal combustors, and six different gasifier systems. Plant sizes ranged from 100 to 1000 MWe. Overall plant performance was calculated using two methodologies. In one task, the channel performance was assumed and the MHD topping cycle efficiencies were based on the assumed values. A second task involved rigorous calculations of channel performance (enthalpy extraction, isentropic efficiency and generator output) that verified the original (task one) assumptions. Closed cycle MHD capital costs were estimated for the task one plants; task two cost estimates were made for the channel and magnet only.

Owens, W.↗
NASA NTRS1976

The role of fracture mechanics in the design of fuel tanks in space vehicles

With special reference to design of fuel tanks in space vehicles, the principles of fracture mechanics are reviewed. An approximate but extremely simple relationship is derived among the operating stress level, the length of crack, and the number of cycles of failure. Any one of the variables can be computed approximately from the knowledge of the other two, if the loading schedule (mission of the tank) is not greatly altered. Two sample examples illustrating the procedures of determining the allowable safe operating stress corresponding to a set of assumed loading schedule are included. The selection of sample examples is limited by the relatively meager available data on the candidate material for various stress ratios in the cycling.

Denton, S. J.↗
NASA NTRS1980

Experimental Study of the Stability of Aircraft Fuels at Elevated Temperatures

An experimental study of fuel stability was conducted in an apparatus which simulated an aircraft gas turbine fuel system. Two fuels were tested: Jet A and Number 2 Home Heating oil. Jet A is an aircraft gas turbine fuel currently in wide use. No. 2HH was selected to represent the properties of future turbine fuels, particularly experimental Reference Broad Specification, which, under NASA sponsorship, was considered as a possible next-generation fuel. Tests were conducted with varying fuel flow rates, delivery pressures and fuel pretreatments (including preheating and deoxygenation). Simulator wall temperatures were varied between 422K and 672K at fuel flows of 0.022 to 0.22 Kg/sec. Coking rate was determined at four equally-spaced locations along the length of the simulator. Fuel samples were collected for infrared analysis. The dependence of coking rate in Jet A may be correlated with surface temperature via an activation energy of 9 to 10 kcal/mole, although the results indicate that both bulk fluid and surface temperature affect the rate of decomposition. As a consequence, flow rate, which controls bulk temperature, must also be considered. Taken together, these results suggest that the decomposition reactions are initiated on the surface and continue in the bulk fluid. The coking rate data for No. 2 HH oil are very highly temperature dependent above approximately 533K. This suggests that bulk phase reactions can become controlling in the formation of coke.

Vranos, A.↗
NASA NTRS1984

Research on aviation fuel instability

The problems associated with aircraft fuel instability are discussed. What is currently known about the problem is reviewed and a research program to identify those areas where more research is needed is discussed. The term fuel instability generally refers to the gums, sediments, or deposits which can form as a result of a set of complex chemical reactions when a fuel is stored for a long period at ambient conditions or when the fuel is thermally stressed inside the fuel system of an aircraft.

Baker, C. E.↗
NASA NTRS1972

Factors affecting altitude relight performance of a double-annular ram-induction combustor

A test program was conducted to evaluate the altitude relight capabilities of a short-length, double-annular, ram-induction combustor which was designed for Mach 3 cruise operation. The use of distorted inlet-air flow profiles was tried to evaluate their effect on the relight performance. No significant improvement in altitude relight performance was obtained with this approach. A study was also made to determine the effects of the reference Mach number, the fuel temperature, and the fuel volatility (ASTM-A1 against JP-4) on the altitude relight performance. Decreasing the reference Mach number, increasing the fuel temperature, and using more volatile fuel all decrease the combustor pressure necessary for relight.

Schultz, D. F.↗
NASA NTRS1980

Experimental evaluation of combustor concepts for burning broad property fuels

A baseline CF6-50 combustor and three advanced combustor designs were evaluated to determine the effects of combustor design on operational characteristics using broad property fuels. Three fuels were used in each test: Jet A, a broad property 13% hydrogen fuel, and a 12% hydrogen fuel blend. Testing was performed in a sector rig at true cruise and simulated takeoff conditions for the CF6-50 engine cycle. The advanced combustors (all double annular, lean dome designs) generally exhibited lower metal temperatures, exhaust emissions, and carbon buildup than the baseline CF6-50 combustor. The sensitivities of emissions and metal temperatures to fuel hydrogen content were also generally lower for the advanced designs. The most promising advanced design used premixing tubes in the main stage. This design was chosen for additional testing in which fuel/air ratio, reference velocity, and fuel flow split were varied.

Kasper, J. M.↗
NASA NTRS1977

Flame characteristics of a NASA contra swirler module

The structure of unconfined flames stabilized on a NASA contra-swirl can has been examined. The structure depends on reference velocity and fuel-air ratio, and photographic illustrations of representative examples of each flame type are shown. A highly compact flame is obtained over a narrow band of fuel-air ratios, within the total range over which flames may be stabilized, at high reference velocity (27.8 m/s). The compact structure can only be obtained with careful fuel nozzle location. Chemical concentration profiles, temperature profiles and sodium chloride seeding experiments all confirm the visual impression of compactness. A flame with 50% greater fuel throughput has a much less compact structure, which is reflected in lower maximum temperature, much more complex chemical concentration distributions and a more diffuse recirculation boundary as observed by sodium chloride seeding than in the compact flame. Each of these flames has not only an evident axial recirculation core but also a toroidal recirculation region over the bluff region between the inner and outer swirlers. In contrast, laser Doppler velocimetry measurements carried out on a flame with lower reference velocity (5.65 m/s) do not indicate the presence of a secondary recirculation region in this flame.

Thompson, D.↗