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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 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

Computing Rotational/Vibrational Dynamics Of Turbine Engines

Turbine Engine Transient Response Analysis program, TETRA 2, developed to calculate dynamic response of structure of turbine engine to applied stimuli. Includes options for calculation of both transient response and steady-state forced response. Helps turbine engineer in understanding basic transient and steady-state responses of whole structure and in identifying crucial problem areas. Could lead to design of turbine engines responding more robustly to applied stimuli. Written in FORTRAN 77.

Gallardo, V. G.↗
NASA NTRS1982

Cost/benefit studies of advanced materials technologies for future aircraft turbine engines: Materials for advanced turbine engines

Cost benefit studies were conducted on six advanced materials and processes technologies applicable to commercial engines planned for production in the 1985 to 1990 time frame. These technologies consisted of thermal barrier coatings for combustor and high pressure turbine airfoils, directionally solidified eutectic high pressure turbine blades, (both cast and fabricated), and mixers, tail cones, and piping made of titanium-aluminum alloys. A fabricated titanium fan blisk, an advanced turbine disk alloy with improved low cycle fatigue life, and a long-life high pressure turbine blade abrasive tip and ceramic shroud system were also analyzed. Technologies showing considerable promise as to benefits, low development costs, and high probability of success were thermal barrier coating, directionally solidified eutectic turbine blades, and abrasive-tip blades/ceramic-shroud turbine systems.

Stearns, M.↗
NASA NTRS1984

Progress in the utilization of an oxide-dispersion-strengthened alloy for small engine turbine blades

The conventional means of improving gas turbine engine performance typically involves increasing the turbine inlet temperature; however, at these higher operational temperatures the high pressure turbine blades require air-cooling to maintain durability. Air-cooling imposes design, material, and economic constraints not only on the turbine blades but also on engine performance. The use of uncooled turbine blades at increased operating temperatures can offer significantly improved performance in small gas turbine engines. A program to demonstrate uncooled MA6000 high pressure turbine blades in a GTEC TFE731 turbofan engine is being conducted. The project goals include demonstration of the advantages of using uncooled MA6000 turbine blades as compared with cast directionally solidified MAR-M 247 blades.

Beatty, T. G.↗
DOE OSTI2024

Water separator for turbine engine

A turbine engine is provided that includes a turbine engine core, a water separate and a core flowpath. The turbine engine core includes a core compressor section, a core combustor section and a core turbine section. The water separator includes a flowpath duct and a collector duct. The flowpath duct extends axially along and circumferentially about an axis. The collector duct extends axially along and circumferentially about the flowpath duct. The collector duct is fluidly coupled with the flowpath duct through a water permeable region in an outer wall of the flowpath duct. The core flowpath extends through the core compressor section, the core combustor section, the core turbine section and the flowpath duct from an inlet into the core flowpath to an exhaust from the core flowpath. The core flowpath extends axially along the water permeable region within the flowpath duct.

Alahyari, Abbas A.↗
NASA NTRS1979

Materials for Advanced Turbine Engines: Power Metallurgy Rene 95 Rotating Turbine Engine Parts - Volume 1

An attempt was made to improve methods for producing powder metallurgy aircraft gas turbine engine parts from the nickel base superalloy known as Rene 95. The parts produced were the high pressure turbine aft shaft for the CF6-50 engine and the stages 5 through 9 compressor disk forgings for the CFM56/F101 engines. A 50% cost reduction was achieved as compared to conventional cast and wrought processing practices. An integrated effort involving several powder producers and a major forging source were included.

Pfouts, W. R.↗
NASA NTRS2016

Gas Turbine Engine Having Fan Rotor Driven by Turbine Exhaust and with a Bypass

A gas turbine engine has a core engine incorporating a core engine turbine. A fan rotor is driven by a fan rotor turbine. The fan rotor turbine is in the path of gases downstream from the core engine turbine. A bypass door is moveable from a closed position at which the gases from the core engine turbine pass over the fan rotor turbine, and moveable to a bypass position at which the gases are directed away from the fan rotor turbine. An aircraft is also disclosed.

Suciu, Gabriel L.↗
NASA NTRSDate not supplied

Transient Optimization of a Gas Turbine Engine

Gas turbine engines are the primary power plants for modern commercial aircraft. Transients prompted by significant changes in thrust or power demand are common and unavoidable. Extreme transient scenarios such as those associated with a go-around during a landing attempt are possible and must be accounted for in the design of the engine and its controller. Engine transients tend to cause a reduction in compressor operability margin, which must be addressed by the engine control system and accounted for in the engine design to prevent events such as compressor stall/surge and combustor blow out. Transient operability concerns typically lead to compromises in the engine design that sacrifice efficiency and/or limit responsiveness. Transient operability is typically managed by logic that limits the fuel flow command. If this logic is not optimized, then the potential for valuable performance could be lost. This study presents a strategy for optimizing the transient limit logic and proposes a strategy for updating the control logic over the lifespan of the engine. The results demonstrate significant improvements in transient operability. For example, of the results at sea level static conditions demonstrated a 31% reduction in the usage of the high pressure compressor operability stack during a snap acceleration transient. Furthermore, a reinforcement learning algorithm is demonstrated to modify the transient logic as the engine degrades to minimize response time while respecting a prescribed compressor operability margin limit. A simple demonstration of the reinforcement learning algorithm resulted in a thrust response time reduction of ~11.8%.

transient↗
NASA NTRSDate not supplied

Transient Optimization of a Gas Turbine Engine

Gas turbine engines are the primary power plants for modern commercial aircraft. Transients prompted by significant changes in thrust or power demand are common and unavoidable. Extreme transient scenarios such as those associated with a go-around during a landing attempt are possible and must be accounted for in the design of the engine and its controller. Engine transients tend to cause a reduction in compressor operability margin, which must be addressed by the engine control system and accounted for in the engine design to prevent events such as compressor stall/surge and combustor blow out. Transient operability concerns typically lead to compromises in the engine design that sacrifice efficiency and/or limit responsiveness. Transient operability is typically managed by logic that limits the fuel flow command. If this logic is not optimized, then the potential for valuable performance could be lost. This study presents a strategy for optimizing the transient limit logic and proposes a strategy for updating the control logic over the lifespan of the engine. The results demonstrate significant improvements in transient operability. For example, of the results at sea level static conditions demonstrated a 31% reduction in the usage of the high pressure compressor operability stack during a snap acceleration transient. Furthermore, a reinforcement learning algorithm is demonstrated to modify the transient logic as the engine degrades to minimize response time while respecting a prescribed compressor operability margin limit. A simple demonstration of the reinforcement learning algorithm resulted in a thrust response time reduction of ~11.8%.

transient↗
NASA NTRSDate not supplied

Transient Optimization of a Gas Turbine Engine

Gas turbine engines are the primary power plants for modern commercial aircraft. Transients prompted by significant changes in thrust or power demand are common and unavoidable. Extreme transient scenarios such as those associated with a go-around during a landing attempt are possible and must be accounted for in the design of the engine and its controller. Engine transients tend to cause a reduction in compressor operability margin, which must be addressed by the engine control system and accounted for in the engine design to prevent events such as compressor stall/surge and combustor blow out. Transient operability concerns typically lead to compromises in the engine design that sacrifice efficiency and/or limit responsiveness. Transient operability is typically managed by logic that limits the fuel flow command. If this logic is not optimized, then the potential for valuable performance could be lost. This study presents a strategy for optimizing the transient limit logic and proposes a strategy for updating the control logic over the lifespan of the engine. The results demonstrate significant improvements in transient operability. For example, of the results at sea level static conditions demonstrated a 31% reduction in the usage of the high pressure compressor operability stack during a snap acceleration transient. Furthermore, a reinforcement learning algorithm is demonstrated to modify the transient logic as the engine degrades to minimize response time while respecting a prescribed compressor operability margin limit. A simple demonstration of the reinforcement learning algorithm resulted in a thrust response time reduction of ~11.8%.

transient↗
NASA NTRS2003

Aircraft engine with inter-turbine engine frame supported counter rotating low pressure turbine rotors

An aircraft gas turbine engine assembly includes an inter-turbine frame axially located between high and low pressure turbines. Low pressure turbine has counter rotating low pressure inner and outer rotors with low pressure inner and outer shafts which are at least in part rotatably disposed co-axially within a high pressure rotor. Inter-turbine frame includes radially spaced apart radially outer first and inner second structural rings disposed co-axially about a centerline and connected by a plurality of circumferentially spaced apart struts. Forward and aft sump members having forward and aft central bores are fixedly joined to axially spaced apart forward and aft portions of the inter-turbine frame. Low pressure inner and outer rotors are rotatably supported by a second turbine frame bearing mounted in aft central bore of aft sump member. A mount for connecting the engine to an aircraft is located on first structural ring.

Seda, Jorge F.↗
DOE OSTI2024

Heat exchanger(s) for recovering water and/or heat energy from turbine engine combustion products

A turbine engine is provided that includes a fan section, a turbine engine core and a recovery system. The turbine engine core is configured to power the fan section. The turbine engine core includes a core compressor section, a core combustor section and a core turbine section. A centerline axis of the turbine engine core is offset from a centerline axis of the turbine engine. The recovery system includes an evaporator module and a condenser module. A core flowpath extends sequentially through the core compressor section, the core combustor section, the core turbine section, the evaporator module and the condenser module from an inlet into the core flowpath to an exhaust from the core flowpath.

Sobanski, Jon E.↗
NASA NTRS2015

Efficient, Low Pressure Ratio Propulsor for Gas Turbine Engines

A gas turbine engine includes a spool, a turbine coupled to drive the spool, and a propulsor that is coupled to be driven by the turbine through the spool. A gear assembly is coupled between the propulsor and the spool such that rotation of the turbine drives the propulsor at a different speed than the spool. The propulsor includes a hub and a row of propulsor blades that extends from the hub. The row includes no more than 20 of the propulsor blades.

Gallagher, Edward J.↗
NASA NTRS2016

Efficient, Low Pressure Ratio Propulsor for Gas Turbine Engines

A gas turbine engine includes a spool, a turbine coupled to drive the spool, and a propulsor that is coupled to be driven by the turbine through the spool. A gear assembly is coupled between the propulsor and the spool such that rotation of the turbine drives the propulsor at a different speed than the spool. The propulsor includes a hub and a row of propulsor blades that extends from the hub. The row includes no more than 20 of the propulsor blades.

Gallagher, Edward J.↗
NASA NTRS1986

Ceramics for turbine engines

The Ceramics for Turbine Engines Project is comprised of three main research programs with major elements as indicated: materials and processing (monolithics and fiber reinforcement), design methodology (design code and tribology), and life prediction (environmental effects, nondestructive evaluation, fracture and fatigue, and time dependent behavior). From the NASA perspective an enhanced ceramics technology base directly supports aeronautics initiatives in small engine technology, high-performance turbine engine technology, and hypersonics. An overview of the program, which includes the technical objectives and content of each program, is provided.

Levine, Stanley R.↗
NASA NTRS1984

Cold-air performance of compressor-drive turbine of department of energy upgraded automobile gas turbine engine. 3: Performance of redesigned turbine

The aerodynamic performance of a redesigned compressor drive turbine of the gas turbine engine is determined in air at nominal inlet conditions of 325 K and 0.8 bar absolute. The turbine is designed with a lower flow factor, higher rotor reaction and a redesigned inlet volute compared to the first turbine. Comparisons between this turbine and the originally designed turbine show about 2.3 percentage points improvement in efficiency at the same rotor tip clearance. Two versions of the same rotor are tested: (1) an as cast rotor, and (2) the same rotor with reduced surface roughness. The effect of reducing surface roughness is about one half percentage point improvement in efficiency. Tests made to determine the effect of Reynolds number on the turbine performance show no effect for the range from 100,000 to 500,000.

Roelke, R. J.↗
NASA NTRS2018

Efficient, Low Pressure Ratio Propulsor for Gas Turbine Engines

A gas turbine engine includes a bypass flow passage that has an inlet and defines a bypass ratio in a range of approximately 8.5 to 13.5. A fan is arranged within the bypass flow passage. A first turbine is a 5-stage turbine and is coupled with a first shaft, which is coupled with the fan. A first compressor is coupled with the first shaft and is a 3-stage compressor. A second turbine is coupled with a second shaft and is a 2-stage turbine. The fan includes a row of fan blades that extend from a hub. The row includes a number (N) of the fan blades, a solidity value (R) at tips of the fab blades, and a ratio of N/R that is from 14 to 16.

Gallagher, Edward J.↗
NASA NTRS2018

Efficient, Low Pressure Ratio Propulsor for Gas Turbine Engines

A gas turbine engine includes a gear assembly and a bypass flow passage that includes an inlet and an outlet that define a design pressure ratio between 1.3 and 1.55. A fan is arranged at the inlet. A first turbine is coupled with a first shaft such that rotation of the first turbine will drive the fan, through the first shaft and the gear assembly, at a lower speed than the first shaft. The fan includes a row of fan blades. The row includes 12-16 (N) fan blades, a solidity value (R) that is from 1.0 to 1.3, and a ratio of N/R that is from 10.0 to 16.

Gallagher, Edward J.↗