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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 379 records · Page 21
All sourcesNASA NTRSDOE OSTI
NASA NTRS2015

Combustion Stability Verification for the Thrust Chamber Assembly of J-2X Developmental Engines 10001, 10002, and 10003

The J-2X engine, a liquid oxygen/liquid hydrogen propellant rocket engine available for future use on the upper stage of the Space Launch System vehicle, has completed testing of three developmental engines at NASA Stennis Space Center. Twenty-one tests of engine E10001 were conducted from June 2011 through September 2012, thirteen tests of the engine E10002 were conducted from February 2013 through September 2013, and twelve tests of engine E10003 were conducted from November 2013 to April 2014. Verification of combustion stability of the thrust chamber assembly was conducted by perturbing each of the three developmental engines. The primary mechanism for combustion stability verification was examining the response caused by an artificial perturbation (bomb) in the main combustion chamber, i.e., dynamic combustion stability rating. No dynamic instabilities were observed in the TCA, although a few conditions were not bombed. Additional requirements, included to guard against spontaneous instability or rough combustion, were also investigated. Under certain conditions, discrete responses were observed in the dynamic pressure data. The discrete responses were of low amplitude and posed minimal risk to safe engine operability. Rough combustion analyses showed that all three engines met requirements for broad-banded frequency oscillations. Start and shutdown transient chug oscillations were also examined to assess the overall stability characteristics, with no major issues observed.

Morgan, C. J.↗
NASA NTRS2019

Control Technology Needs for Electrified Aircraft Propulsion Systems

Electrified aircraft propulsion (EAP) systems hold potential for the reduction of aircraft fuel burn, emissions, and noise. Currently, NASA and other organizations are actively working to identify and mature technologies necessary to bring EAP designs to reality. This paper specifically focuses on the envisioned control technology challenges associated with EAP designs that include gas turbine technology. Topics discussed include analytical tools for the dynamic modeling and analysis of EAP systems, and control design strategies at the propulsion and component levels. This includes integrated supervisory control facilitating the coordinated operation of turbine and electrical components, control strategies that seek to minimize fuel consumption and lessen the challenges associated with thermal management, and dynamic control to ensure engine operability during system transients. These dynamic control strategies include innovative control approaches that either extract or supply power to engine shafts dependent upon operating phase, which may improve performance and reduced gas turbine engine weight. Finally, a discussion of control architecture design considerations to help alleviate the propulsion/aircraft integration and certification challenges associated with EAP systems is provided.

Simon, Donald L.↗
NASA NTRS2019

Control Technology Needs for Electrified Aircraft Propulsion Systems

Electrified aircraft propulsion (EAP) systems hold potential for the reduction of aircraft fuel burn, emissions, and noise. Currently, NASA and other organizations are actively working to identify and mature technologies necessary to bring EAP designs to reality. This paper specifically focuses on the envisioned control technology challenges associated with EAP designs that include gas turbine technology. Topics discussed include analytical tools for the dynamic modeling and analysis of EAP systems, and control design strategies at the propulsion and component levels. This includes integrated supervisory control facilitating the coordinated operation of turbine and electrical components, control strategies that seek to minimize fuel consumption and lessen the challenges associated with thermal management, and dynamic control to ensure engine operability during system transients. These dynamic control strategies include innovative control approaches that either extract or supply power to engine shafts dependent upon operating phase, which may improve performance and reduced gas turbine engine weight. Finally, a discussion of control architecture design considerations to help alleviate the propulsion/aircraft integration and certification challenges associated with EAP systems is provided.

Engine Controls↗
NASA NTRS2019

Control Technology Needs for Electrified Aircraft Propulsion Systems

Electrified aircraft propulsion (EAP) systems hold potential for the reduction of aircraft fuel burn, emissions, and noise. Currently, NASA and other organizations are actively working to identify and mature technologies necessary to bring EAP designs to reality. This paper specifically focuses on the envisioned control technology challenges associated with EAP designs that include gas turbine technology. Topics discussed include analytical tools for the dynamic modeling and analysis of EAP systems, and control design strategies at the propulsion and component levels. This includes integrated supervisory control facilitating the coordinated operation of turbine and electrical components, control strategies that seek to minimize fuel consumption and lessen the challenges associated with thermal management, and dynamic control to ensure engine operability during system transients. These dynamic control strategies include innovative control approaches that either extract or supply power to engine shafts dependent upon operating phase, which may improve performance and reduced gas turbine engine weight. Finally, a discussion of control architecture design considerations to help alleviate the propulsion/aircraft integration and certification challenges associated with EAP systems is provided.

Gas Turbine Engines↗
NASA NTRS2013

A Model to Assess the Risk of Ice Accretion Due to Ice Crystal Ingestion in a Turbofan Engine and its Effects on Performance

The occurrence of ice accretion within commercial high bypass aircraft turbine engines has been reported under certain atmospheric conditions. Engine anomalies have taken place at high altitudes that were attributed to ice crystal ingestion, partially melting, and ice accretion on the compression system components. The result was one or more of the following anomalies: degraded engine performance, engine roll back, compressor surge and stall, and flameout of the combustor. The main focus of this research is the development of a computational tool that can estimate whether there is a risk of ice accretion by tracking key parameters through the compression system blade rows at all engine operating points within the flight trajectory. The tool has an engine system thermodynamic cycle code, coupled with a compressor flow analysis code, and an ice particle melt code that has the capability of determining the rate of sublimation, melting, and evaporation through the compressor blade rows. Assumptions are made to predict the complex physics involved in engine icing. Specifically, the code does not directly estimate ice accretion and does not have models for particle breakup or erosion. Two key parameters have been suggested as conditions that must be met at the same location for ice accretion to occur: the local wet-bulb temperature to be near freezing or below and the local melt ratio must be above 10%. These parameters were deduced from analyzing laboratory icing test data and are the criteria used to predict the possibility of ice accretion within an engine including the specific blade row where it could occur. Once the possibility of accretion is determined from these parameters, the degree of blockage due to ice accretion on the local stator vane can be estimated from an empirical model of ice growth rate and time spent at that operating point in the flight trajectory. The computational tool can be used to assess specific turbine engines to their susceptibility to ice accretion in an ice crystal environment.

Jorgenson, Philip C. E.↗
NASA NTRS1975

Risk Management Technique for design and operation of facilities and equipment

The Risk Management System collects information from engineering, operating, and management personnel to identify potentially hazardous conditions. This information is used in risk analysis, problem resolution, and contingency planning. The resulting hazard accountability system enables management to monitor all identified hazards. Data from this system are examined in project reviews so that management can decide to eliminate or accept these risks. This technique is particularly effective in improving the management of risks in large, complex, high-energy facilities. These improvements are needed for increased cooperation among industry, regulatory agencies, and the public.

Fedor, O. H.↗
DOE OSTI2021

Advancing Pressure Gain Combustion for Terrestrial Turbine Systems

The principal objectives of this project were to demonstrate the operability of a natural gas - air RDE at gas turbine engine cycle conditions. To achieve these goals, our specific tasks were to 1) characterize the dynamic injection and mixing performance of candidate injectors, 2) assess the engine operability via interrogation of spatio-temporal dynamics of combustor inlet and exit flows, 3) measure the engine emissions, and 4) complete an overall assessment of the pressure gain potential of the device. In order to advance RDE technology to the greatest extent possible, the reactants and experimental conditions were chosen to replicate combustor and injector conditions in a terrestrial power generation system. This path was taken because the scaling of combustion devices is, in general, an immature area given the challenges of interrelated mixing and chemical kinetics processes both of which scale with operating conditions.

03 NATURAL GAS↗
DOE OSTI2021

I.1 Heavy-Duty Diesel Combustion (Sandia National Laboratories)

Regulatory drivers and market demands for lower pollutant emissions, lower carbon dioxide emissions, and lower fuel consumption motivate the development of cleaner and more fuel-efficient engine operating strategies. Most current production heavy-duty diesel engines use a combination of both in-cylinder and exhaust emissions-control strategies to achieve these goals. The emissions and efficiency performance of in-cylinder strategies depend strongly on flow and mixing processes that can be influenced by using multiple fuel injections. Past work performed under this project showed that adding a second injection can reduce soot to levels below what would have been produced by an unchanged first injection, thereby increasing load while decreasing soot and potentially reducing brake specific fuel consumption. Information characterizing the important in-cylinder processes with multiple injections has been gleaned from ensemble-averaged planar laser-induced incandescence (PLII) imaging visualizing the soot cloud and planar induced fluorescence (PLIF) of OH characterizing the soot oxidation regions. PLII showed a consistent disruption of the first injection soot cloud by the second injection. In conjunction with OH-PLIF, differences in soot oxidation patterns for multiple injections compared to single injections were observed. This understanding was further enhanced in FY20, when high-speed imaging resolving the above-mentioned effects in a single cycle were combined with direct numerical simulations investigating the multiple-injection ignition process on the microscopic level of turbulence and chemistry interaction. In FY21, these findings in conjunction with findings from other researchers published in the scientific literature were composed into a preliminary multiple-injection conceptual model of fuel-mixing, injection and ignition processes. Remaining key research questions were also highlighted. In addition, wall heat flux was investigated experimentally and with numerical simulations to understand the potential of multiple injections to reduce the engine heat losses and further enhance the efficiency.

33 ADVANCED PROPULSION SYSTEMS↗
NASA NTRS1978

Recent advances in convectively cooled engine and airframe structures for hypersonic flight

Research over the past decade has identified critical thermal/structural design problems and has produced viable design concepts for a second generation experimental scramjet. The design concepts for the hydrogen fuel-cooled engine structure involve a variety of innovative features to accommodate the harsh aerothermal environment encountered within the engine. The baseline concept which has evolved has reasonable mass characteristics, and cooling requirements which permit engine operation to Mach 9-10 without additional hydrogen for engine cooling. Studies have identified fabrication techniques and coolant passage configurations that increase fatigue life of the structure an order of magnitude over previous configurations.

Kelly, H. N.↗
NASA NTRS1979

Application of digital controls on the quiet clean short haul experimental engines

The digital control systems for the Under-the-Wing (UTW) and Over-the-Wing (OTW) engines developed on the NASA/General Electric Quiet Clean Shorthaul Experimental Engine (QCSEE) program are described. The system to control engine variables includes three major functional parts: system sensors, digital control and system actuators. One of the primary control system functions is to prevent the engine from exceeding speed or temperature limits. The UTW control system also provides fault detection and condition monitoring. The control system requirements for the OTW engine are essentially the same as the UTW engine, however the inlet Mach number control requirement is eliminated, and failure indication and corrective action and full authority digital control are added. The digital controls scheduled the engine variables and maintained engine operation within all physical limits throughout the test program of approximately 200 hours of operation and provided stable and accurate control of both engines.

Saunders, A. A., Jr.↗
NASA NTRS1951

NACA Investigations of Icing-Protection Systems for Turbojet-Engine Installations

Investigations have been made in flight and in wind tunnels to determine which components of turbojet installations are most critical in icing conditions, and to evaluate several methods of icing protection. From these studies, the requirements necessary for adequate icing protection and the consequent penalties on engine performance can be estimated. Because investigations have indicated that the compressor-inlet screen constitutes the greatest icing hazard and is difficult to protect, complete removal or retraction of the screen upon encountering an icing condition is recommended. In the absence of the screen, the inlet guide vanes of an axial-flow-type turbojet engine constitute the greatest danger to engine operation in an icing condition; a centrifugal-type engine, on the other hand, is relatively unsusceptible to icing once the screen has been removed. Of the three icing-protection systems investigated, surface heating, hot-gas bleedback, and inertia-separation inlets, only the first two offer an acceptable solution to the problem of engine icing protection. Surface heating, either by gas heating or electrical means, appears to be the most acceptable icing-protection method with regard to performance losses. Hot-gas bleedback, although causing undesirable thrust losses, offers an easy means of obtaining icing protection for some installations. The final choice of an icing-protection system depends, however, on the supply of heated gas and electrical power available and on the allowable performance and. weight penalties associated with each system.

VonGlahn, Uwe↗
NASA NTRS1990

Thermal barrier coating evaluation needs

A 0.025 cm (0.010 in) thick thermal barrier coating (TBC) applied to turbine airfoils in a research gas turbine engine provided component temperature reductions of up to 190 C. These impressive temperature reductions can allow increased engine operating temperatures and reduced component cooling to achieve greater engine performance without sacrificing component durability. The significant benefits of TBCs are well established in aircraft gas turbine engine applications and their use is increasing. TBCs are also under intense development for use in the Low Heat Rejection (LHR) diesel engine currently being developed and are under consideration for use in utility and marine gas turbines. However, to fully utilize the benefits of TBCs it is necessary to accurately characterize coating attributes that affect the insulation and coating durability. The purpose there is to discuss areas in which nondestructive evaluation can make significant contributions to the further development and full utilization of TBCs for aircraft gas turbine engines and low heat rejection diesel engines.

Brindley, William J.↗
NASA NTRS1991

Control center operations at the Goddard Space Flight Center

Information is given in viewgraph form on control center operations at the Goddard Space Flight Center. Information is given on multi-mission, dedicated and attached payload support; the mission operations environment; technical challenges; Hubble Space Telescope (HST) science planning interface; operations engineering life cycle; Special Analysis of Maneuverable Satellites (SAMS) operations; and user interface applications.

Connerton, Robert M.↗
NASA NTRS1992

Ground signature extrapolation of three-dimensional near-field CFD predictions for several HSCT configurations

A CFD analysis of the near-field sonic boom environment of several low boom High Speed Civilian Transport (HSCT) concepts is presented. The CFD method utilizes a multi-block Euler marching code within the context of an innovative mesh topology that allows for the resolution of shock waves several body lengths from the aircraft. Three-dimensional pressure footprints at one body length below three-different low boom aircraft concepts are presented. Models of two concepts designed by NASA to cruise at Mach 2 and Mach 3 were built and tested in the wind tunnel. The third concept was designed by Boeing to cruise at Mach 1.7. Centerline and sideline samples of these footprints are then extrapolated to the ground using a linear waveform parameter method to estimate the ground signatures or sonic boom ground overpressure levels. The Mach 2 concept achieved its centerline design signature but indicated higher sideline booms due to the outboard wing crank of the configuration. Nacelles are also included on two of NASA's low boom concepts. Computations are carried out for both flow-through nacelles and nacelles with engine exhaust simulation. The flow-through nacelles with the assumption of zero spillage and zero inlet lip radius showed very little effect on the sonic boom signatures. On the other hand, it was shown that the engine exhaust plumes can have an effect on the levels of overpressure reaching the ground depending on the engine operating conditions. The results of this study indicate that engine integration into a low boom design should be given some attention.

Siclari, M. J.↗
NASA NTRS2004

Furnace Cyclic Oxidation Behavior of Multi-Component Low Conductivity Thermal Barrier Coatings

Ceramic thermal barrier coatings will play an increasingly important role in advanced gas turbine engines because of their ability to further increase engine operating temperatures and reduce cooling, thus helping achieve future engine low emission, high efficiency and improved reliability goals. Advanced multi-component zirconia-based thermal barrier coatings are being developed using an oxide defect clustering design approach to achieve the required coating low thermal conductivity and high temperature stability. Although the new composition coatings were not yet optimized for cyclic durability, an initial durability screening of the candidate coating materials was conducted using conventional furnace cyclic oxidation tests. In this paper, furnace cyclic oxidation behavior of plasma-sprayed zirconia-based defect cluster thermal barrier coatings was investigated at 1163 C using 45 min hot cycles. The ceramic coating failure mechanisms were studied using scanning electron microscopy (SEM) combined with X-ray diffraction (XRD) phase analysis after the furnace tests. The coating cyclic lifetime is also discussed in relation to coating processing, phase structures, dopant concentration, and other thermo-physical properties.

Zhu, Dong-Ming↗
NASA NTRS2005

Thermal and Environmental Barrier Coatings for Advanced Turbine Engine Applications

Ceramic thermal and environmental barrier coatings (T/EBCs) will play a crucial role in advanced gas turbine engine systems because of their ability to significantly increase engine operating temperatures and reduce cooling requirements, thus help achieve engine low emission and high efficiency goals. Advanced T/EBCs are being developed for the low emission SiC/SiC ceramic matrix composite (CMC) combustor applications by extending the CMC liner and vane temperature capability to 1650 C (3000 F) in oxidizing and water vapor containing combustion environments. Low conductivity thermal barrier coatings (TBCs) are also being developed for metallic turbine airfoil and combustor applications, providing the component temperature capability up to 1650 C (3000 F). In this paper, ceramic coating development considerations and requirements for both the ceramic and metallic components will be described for engine high temperature and high-heat-flux applications. The underlying coating failure mechanisms and life prediction approaches will be discussed based on the simulated engine tests and fracture mechanics modeling results.

Zhu, Dong-Ming↗
NASA NTRS2003

Therma1 Conductivity and Durability of Advanced Thermal Barrier Coatings

Thermal barrier coatings (TBCs) will play a crucial role in advanced gas turbine engines because of their ability to further increase engine operating temperature and reduce cooling, thus helping to achieve engine emission and efficiency goals. Future TBCs must be designed with increased phase stability, lower thermal conductivity, and improved sintering and thermal stress resistance in order to effectively protect engine hot-section components. Advanced low conductivity TBCs are being developed at NASA by incorporating multi-component oxide dopants into zirconia-yttria or hafnia-yttria to promote the formation of thermodynamically stable defect clusters within the coating structures. This presentation will primarily focus on thermal conductivity and durability of the novel defect cluster thermal barrier coatings for turbine airfoil and combustor applications, determined by a unique CO2 laser heat-flux approach. The laser heat-flux testing approach emphasizes the real-time monitoring and assessment of the coating thermal conductivity under simulated engine temperature and thermal gradient conditions. The conductivity increase due to coating sintering (and/or phase change) and the conductivity decrease due to coating delamination have been determined under steady-state, cyclic, uniform or non-uniform heat-flux conditions. The coating radiation flux resistance has been evaluated by varying coating thermal gradients, and also by using a laser-heated radiative-flux source. Advanced multi-component TBC systems have been shown to have significantly reduced thermal conductivity and improved high temperature stability due to the nano-sized, low mobility defect clusters associated with the paired rare earth dopant additions. The effect of oxide defect cluster dopants on coating thermal conductivity, thermal stability and furnace cyclic durability will also be discussed. The current low conductivity TBC systems have demonstrated long-term cyclic durability at very high temperatures which is far beyond the baseline ZrO2-Y2O3 capabilities.

Zhu, Dong-Ming↗