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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

Combustor technology for future aircraft

The continuing improvement of aircraft gas turbine engine operating efficiencies involves increases in overall engine pressure ratio increases that will result in combustor inlet pressure and temperature increases, greater combustion temperature rises, and higher combustor exit temperatures. These conditions entail the development of fuel injectors generating uniform circumferential and radial temperature patterns, as well as combustor liner configurations and materials capable of withstanding increased thermal radiation even as the amount of cooling air is reduced. Low NO(x)-emitting combustor concepts are required which will employ staged combustion. The development status of component technologies answering these requirements are presently evaluated.

Tacina, Robert R.↗

Performance Characteristics of a Normal-shock Side Inlet Located Downstream of a Canard Control Surface at Mach Numbers of 1.5 and 1.8/

The performance characteristics of a downward canted normal-shock side (scoop) inlet located downstream of a triangular control surface are presented for free-stream Mach numbers of 1.5 and 1.8 in terms of total pressure recovery and mass flow ratio for various boundary-layer removal systems,angles of attack, control surface deflections and adverse yaw. An engine operating condition for a hypothetical turbojet engine is established, and the match point characteristics of the engine-inlet configuration are summarized. 520::It is shown that the diffuser performance increases with increased boundary-layer removal and decreases because of the presence of the wake from the forward control surface. At the higher angles of attack the wake passes over the inlet and does not affect the inlet performance. Adverse yaw reduces the total pressure recovery values below those for the unawed case. Magnitudes of the total pressure recovery were below the theoretical normal-shock recovery for the respective test Mach numbers.

INLETS, SIDE - DUCTED BODIES↗

Temperature Dependent Modal Test/Analysis Correlation of X-34 Fastrac Composite Rocket Nozzle

A unique high temperature modal test and model correlation/update program has been performed on the composite nozzle of the FASTRAC engine for the NASA X34 Reusable Launch Vehicle. The program was required to provide an accurate high temperature model of nozzle for incorporation into the engine system structural dynamics model for loads calculation; this model is significantly different from the ambient case due to the large decrease in composite material properties due to heating. The modal test was enabled by piggybacking onto a hot-fire test of the nozzle at NASA/Marshall and tracking the natural frequencies as they decreased. A series of high fidelity modal tests and finite element model correlation of the nozzle in a free-free configuration was initially performed. This model was then attached to a modal-test verified model of the engine hot-fire test stand and the' ambient system mode shapes identified. A reduced set of accelerometers was then attached to the nozzle, the engine fired full-duration, and the frequency peaks corresponding to the ambient nozzle modes individually isolated and tracked. To update the finite element model of the nozzle to these frequency curves, a multiplicative factor was. applied to the rate of decline of the composite material property versus temperature table. This new property table was used to create high-temperature nozzle models corresponding to 10 second engine operation increments and tied into the engine system model for loads determination.

Brown, Andrew M.↗

Temperature Dependent Modal Test/Analysis Correlation of X-34 Fastrac Composite Rocket Nozzle

A unique high temperature modal test and model correlation/update program has been performed on the composite nozzle of the FASTRAC engine for the NASA X-34 Reusable Launch Vehicle. The program was required to provide an accurate high temperature model of the nozzle for incorporation into the engine system structural dynamics model for loads calculation; this model is significantly different from the ambient case due to the large decrease in composite stiffness properties due to heating. The high-temperature modal test was performed during a hot-fire test of the nozzle. Previously, a series of high fidelity modal tests and finite element model correlation of the nozzle in a free-free configuration had been performed. This model was then attached to a modal-test verified model of the engine hot-fire test stand and the ambient system mode shapes were identified. A reduced set of accelerometers was then attached to the nozzle, the engine fired full-duration, and the frequency peaks corresponding to the ambient nozzle modes individually isolated and tracked as they decreased during the test. To update the finite-element model of the nozzle to these frequency curves, the percentage differences of the anisotropic composite moduli due to temperature variation from ambient, which had been used in the initial modeling and which were obtained by small sample coupon testing, were multiplied by an iteratively determined constant factor. These new properties were used to create high-temperature nozzle models corresponding to 10 second engine operation increments and tied into the engine system model for loads determination.

Brown, Andrew M.↗

Utilizing a Common Model Architecture for the Simulation of On-Orbit Human Spaceflight Operations

Advances in computing and simulation technology during the 1980 s promoted the development of multiple simulations for on-orbit human spaceflight operations at the National Aeronautics and Space Administration (NASA) Johnson Space Center (JSC). By the late 1980 s, it became increasingly clear that redundancy risked leading to inefficiencies in dissemination and incorporation of models across simulations, institutional organizations, and spaceflight programs. Software development tools, such as the Trick Simulation Environment, began to emerge providing a generalized framework for the development, integration, and operation of simulations. Throughout the 1990s, engineering, operations, and training simulations began to migrate to Trick in order to take advantage of a common simulation environment. Due to forecasted reductions in funding to the Shuttle and International Space Station (ISS) programs in the early 2000s, the common framework that arose from this simulation migration was formalized as the Common Model Architecture (CMA). The CMA was established in order to further reduce redundancy, while promoting model sharing and enhancing model integrity. A minimal set of standards for data flows, functional interactions, and model organization based on Trick were defined for those features found to be most important in facilitating model interaction. Model-providing organizations within NASA JSC took ownership of models for their specific areas of expertise, and agreements were reached on common model sharing and usage. This paper describes the evolution of commonality enabled by the Trick Simulation Environment at JSC leading to the CMA, the standards defined by the CMA, a recent CMA simulation example, and potential application of the CMA to NASA s new Exploration program.

Leslie J Quiocho↗

Demonstration of Active Combustion Control

The primary objective of this effort was to demonstrate active control of combustion instabilities in a direct-injection gas turbine combustor that accurately simulates engine operating conditions and reproduces an engine-type instability. This report documents the second phase of a two-phase effort. The first phase involved the analysis of an instability observed in a developmental aeroengine and the design of a single-nozzle test rig to replicate that phenomenon. This was successfully completed in 2001 and is documented in the Phase I report. This second phase was directed toward demonstration of active control strategies to mitigate this instability and thereby demonstrate the viability of active control for aircraft engine combustors. This involved development of high-speed actuator technology, testing and analysis of how the actuation system was integrated with the combustion system, control algorithm development, and demonstration testing in the single-nozzle test rig. A 30 percent reduction in the amplitude of the high-frequency (570 Hz) instability was achieved using actuation systems and control algorithms developed within this effort. Even larger reductions were shown with a low-frequency (270 Hz) instability. This represents a unique achievement in the development and practical demonstration of active combustion control systems for gas turbine applications.

Lovett, Jeffrey A.↗

Development and Performance Evaluations of HfO2-Si and Rare Earth-Si Based Environmental Barrier Bond Coat Systems for SiC/SiC Ceramic Matrix Composites

Ceramic environmental barrier coatings (EBC) and SiCSiC ceramic matrix composites (CMCs) will play a crucial role in future aircraft propulsion systems because of their ability to significantly increase engine operating temperatures, improve component durability, reduce engine weight and cooling requirements. Advanced EBC systems for SiCSiC CMC turbine and combustor hot section components are currently being developed to meet future turbine engine emission and performance goals. One of the significant material development challenges for the high temperature CMC components is to develop prime-reliant, high strength and high temperature capable environmental barrier coating bond coat systems, since the current silicon bond coat cannot meet the advanced EBC-CMC temperature and stability requirements. In this paper, advanced NASA HfO2-Si based EBC bond coat systems for SiCSiC CMC combustor and turbine airfoil applications are investigated. The coating design approach and stability requirements are specifically emphasized, with the development and implementation focusing on Plasma Sprayed (PS) and Electron Beam-Physic Vapor Deposited (EB-PVD) coating systems and the composition optimizations. High temperature properties of the HfO2-Si based bond coat systems, including the strength, fracture toughness, creep resistance, and oxidation resistance were evaluated in the temperature range of 1200 to 1500 C. Thermal gradient heat flux low cycle fatigue and furnace cyclic oxidation durability tests were also performed at temperatures up to 1500 C. The coating strength improvements, degradation and failure modes of the environmental barrier coating bond coat systems on SiCSiC CMCs tested in simulated stress-environment interactions are briefly discussed and supported by modeling. The performance enhancements of the HfO2-Si bond coat systems with rare earth element dopants and rare earth-silicon based bond coats are also highlighted. The advanced bond coat systems, when integrated with advanced EBC top coats, showed promise to achieve 1500 C temperature capability, helping enable next generation turbine engines with significantly improved engine component temperature capability and long-term durability.

HfO2-Si bond coat↗

Development and Property Evaluation of Selected HfO2-Silicon and Rare Earth-Silicon Based Bond Coats and Environmental Barrier Coating Systems for SiC/SiC Ceramic Matrix Composites

Ceramic environmental barrier coatings (EBC) and SiC/SiC ceramic matrix composites (CMCs) will play a crucial role in future aircraft propulsion systems because of their ability to significantly increase engine operating temperatures, improve component durability, reduce engine weight and cooling requirements. Advanced EBC systems for SiC/SiC CMC turbine and combustor hot section components are currently being developed to meet future turbine engine emission and performance goals. One of the significant material development challenges for the high temperature CMC components is to develop prime-reliant, high strength and high temperature capable environmental barrier coating bond coat systems, since the current silicon bond coat cannot meet the advanced EBC-CMC temperature and stability requirements. In this paper, advanced NASA HfO2-Si and rare earth Si based EBC bond coat EBC systems for SiC/SiC CMC combustor and turbine airfoil applications are investigated. High temperature properties of the advanced EBC systems, including the strength, fracture toughness, creep and oxidation resistance have been studied and summarized. The advanced NASA EBC systems showed some promise to achieve 1500C temperature capability, helping enable next generation turbine engines with significantly improved engine component temperature capability and durability.

Oxidation and cyclic durability↗

Environmental Stability and Oxidation Behavior of HfO2-Si and YbGd(O) Based Environmental Barrier Coating Systems for SiCSiC Ceramic Matrix Composites

Ceramic environmental barrier coatings (EBC) and SiCSiC ceramic matrix composites (CMCs) will play a crucial role in future aircraft propulsion systems because of their ability to significantly increase engine operating temperatures, improve component durability, reduce engine weight and cooling requirements. Advanced EBC systems for SiCSiC CMC turbine and combustor hot section components are currently being developed to meet future turbine engine emission and performance goals. One of the significant material development challenges for the high temperature CMC components is to develop prime-reliant, environmental durable environmental barrier coating systems. In this paper, the durability and performance of advanced Electron Beam-Physical Vapor Deposition (EB-PVD) NASA HfO2-Si and YbGdSi(O) EBC bond coat top coat systems for SiCSiC CMC have been summarized. The high temperature thermomechanical creep, fatigue and oxidation resistance have been investigated in the laboratory simulated high-heat-flux environmental test conditions. The advanced NASA EBC systems showed promise to achieve 1500C temperature capability, helping enable next generation turbine engines with significantly improved engine component temperature capability and durability.

Ceramic Matrix Composites↗

Core/Combustor-Noise Measurements and Source Separation for the DGEN Aeropropulsion Research Turbofan

Acoustic data obtained using a small turbofan engine are further analyzed with the aim to ultimately enable improved prediction methods and mitigation techniques for turbofan core/combustor noise. The relative impact of this propulsion-noise source for civilian-transport aircraft on airport community noise is expected to significantly increase in the future. Simultaneous high-data-rate acoustic measurements acquired using a circumferential infinite-tube-probe array at the core-nozzle exit in conjunction with a far-field microphone array are processed. The test matrix contains engine operational points from idle to engine-maximum setting, with measurements repeated for different circumferential array configurations, as well as for redundancy. The combustor-noise contribution to the far-field noise signature is obtained using an advanced source-separation method that correlates far-field microphone measurements with a modal decomposition of the unsteady pressure field at the core-nozzle exit. The advantages of the present approach compared to the classical two-signal source-separation method are discussed.

Aeroacoustics↗

Preliminary Altitude Operational Characteristics of a J57-P1 Turbojet Engine

The operational characteristics of a J57-P1 turbojet engine have been investigated at altitudes between 15,000 and 66,000 feet in the Lewis altitude wind tunnel. Included in this study is a discussion of fuel nozzle coking, the altitude operating limits with and without the standard engine control, the compressor surge characteristics, and the engine starting and windmilling characteristics. Severe circumferential turbine outlet temperature gradients which occurred at high altitude as a result of fuel nozzle coking were alleviated by the manufacturer's change in the fuel flow divider schedule and in a nozzle gasket material. Compressor air bleed is required to prevent surge of the outboard compressor in the low engine speed region. The maximum altitude at which the engine was operated without the control was about 66,000 feet at 0.8 flight Mach number and at a reduced engine speed to avoid compressor surge; with the engine control in operation, the altitude operating limit is reduced to approximately 59,000 feet. The maximum altitude at which the engine was started was about 40,000 feet.

Wallner, Lewis E.↗

A Systems-Level Perspective on Engine Ice Accretion

The accretion of ice in the compression system of commercial gas turbine engines operating in high ice water content conditions is a safety issue being studied by the aviation sector. While most of the research focuses on the underlying physics of ice accretion and the meteorological conditions in which accretion can occur, a systems-level perspective on the topic lends itself to potential near-term operational improvements. This work focuses on developing an accurate and reliable algorithm for detecting the accretion of ice in the low pressure compressor of a generic 40,000 lbf thrust class engine. The algorithm uses only the two shaft speed sensors and works regardless of engine age, operating condition, and power level. In a 10,000-case Monte Carlo simulation, the detection approach was found to have excellent capability at determining ice accretion from sensor noise with detection occurring when ice blocks an average of 6.8% of the low pressure compressor area. Finally, an initial study highlights a potential mitigation strategy that uses the existing engine actuators to raise the temperature in the low pressure compressor in an effort to reduce the rate at which ice accretes.

May, Ryan D.↗

An Approach to Detect and Mitigate Ice Particle Accretion in Aircraft Engine Compression Systems

The accretion of ice in the compression system of commercial gas turbine engines operating in high ice water content conditions is a safety issue being studied by the aviation sector. While most of the research focuses on the underlying physics of ice accretion and the meteorological conditions in which accretion can occur, a systems-level perspective on the topic lends itself to potential near-term operational improvements. This work focuses on developing an accurate and reliable algorithm for detecting the accretion of ice in the low pressure compressor of a generic 40,000 lbf thrust class engine. The algorithm uses only the two shaft speed sensors and works regardless of engine age, operating condition, and power level. In a 10,000-case Monte Carlo simulation, the detection approach was found to have excellent capability at determining ice accretion from sensor noise with detection occurring when ice blocks an average of 6.8% of the low pressure compressor area. Finally, an initial study highlights a potential mitigation strategy that uses the existing engine actuators to raise the temperature in the low pressure compressor in an effort to reduce the rate at which ice accretes.

May, Ryan D.↗

An Approach to Detect and Mitigate Ice Particle Accretion in Aircraft Engine Compression Systems

The accretion of ice in the compression system of commercial gas turbine engines operating in high ice water content conditions is a safety issue being studied by the aviation sector. While most of the research focuses on the underlying physics of ice accretion and the meteorological conditions in which accretion can occur, a systems-level perspective on the topic lends itself to potential near-term operational improvements. This work focuses on developing an accurate and reliable algorithm for detecting the accretion of ice in the low pressure compressor of a generic 40,000 lbf thrust class engine. The algorithm uses only the two shaft speed sensors and works regardless of engine age, operating condition, and power level. In a 10,000-case Monte Carlo simulation, the detection approach was found to have excellent capability at determining ice accretion from sensor noise with detection occurring when ice blocks an average of 6.8 percent of the low pressure compressor area. Finally, an initial study highlights a potential mitigation strategy that uses the existing engine actuators to raise the temperature in the low pressure compressor in an effort to reduce the rate at which ice accretes.

May, Ryan D.↗

Review of jet engine emissions

A review of the emission characteristics of jet engines is presented. The sources and concentrations of the various constituents in the engine exhaust and the influence of engine operating conditions on emissions are discussed. Cruise emissions to be expected from supersonic engines are compared with emissions from subsonic engines. The basic operating principles of the gas turbine combustor are reviewed together with the effects of combustor operating conditions on emissions. The performance criteria that determine the design of gas turbine combustors are discussed. Combustor design techniques are considered that may be used to reduce emissions.

Grobman, J. S.↗

Combustion-chamber Performance Characteristics of a Python Turbine-propeller Engine Investigated in Altitude Wind Tunnel

Combustion-chamber performance characteristics of a Python turbine-propeller engine were determined from investigation of a complete engine over a range of engine speeds and shaft horsepowers at simulated altitudes. Results indicated the effect of engine operating conditions and altitude on combustion efficiency and combustion-chamber total pressure losses. Performance of this vaporizing type combustion chamber was also compared with several atomizing type combustion chambers. Over the range of test conditions investigated, combustion efficiency varied from approximately 0.95 to 0.99.

COMBUSTION - EFFECT OF ENGINE OPERATING CONDITIONS↗

Preliminary base heating environments for a generalized ALS LO2/LH2 launch vehicle, appendix 1 and 2

A secondary objective of contract NAS8-39141 is to provide base heating assessments, as required, to support Advanced Launch System (ALS) preliminary launch vehicle and propulsion system design studies. The ALS propulsion systems integration working group meeting (No. 3) recently completed in San Diego, California, focused attention on the need for base heating environment determination to provide preliminary requirements for LO2/LH2 propulsion systems currently being considered for ALS. We were requested to provide these environments for a range of possible propellant mixture and nozzle area ratios. Base heating environments can only be determined as a function of altitude when the engine operating conditions and vehicle base region geometry (engine arrangement) are known. If time dependent environments are needed to assess thermal loads, a trajectory must also be provided. These parameters are not fixed at this time since the ALS configurations and propulsion operating conditions are varied and continue to be studied by Phase B contractors. Therefore, for this study, a generalized LO2/LH2 system was selected along with a vehicle configuration consisting of a seven-engine booster and a three-engine core. MSFC provided guidance for the selection. We also selected a limited number of body points on the booster and core vehicles and engines for the environment estimates. Environments at these locations are representative of maximum heating conditions in the base region and are provided as a function of altitude only. Guidelines and assumptions for this assessment, methodology for determining the environments, and preliminary results are provided in this technical note. Refinements in the environments will be provided as the ALS design matures.

Bender, Robert L.↗

Spiral-Groove Ring Seal for Counterrotating Shafts

Self-lubricating seal tolerates high sliding speeds. Application of self-acting geometry in form of spiral grooves to faces of ring-seal housing maintains thin air film of relatively high stiffness between seal ring and housing, enabling seal to operate in noncontacting mode over entire engine-operating range. Potential application in sealing fan-bleed air between two counter-rotating shafts in advanced gas-turbine engines.

Dirusso, E.↗