SCIENTIFIC AND TECHNICAL AEROSPACE REPORTS
A Biweekly Publication of the National Aeronautics and Space Administration
VOLUME 43, ISSUE 20 - OCTOBER 07, 2005
07 AIRCRAFT PROPULSION AND POWER
Includes primary propulsion systems and related systems and components, e.g., gas turbine engines, compressors, and fuel systems; and onboard auxiliary power plants for aircraft.
For related information see also 20 Spacecraft Propulsion and Power; 28 Propellants and Fuels; and 44 Energy Production and Conversion.
20050214846 NASA Glenn Research Center, Cleveland, OH, USA
Glenn-HT Code Validated for Complex Turbine Blade Cooling Passage
Rigby, David L.; Research and Technology 2002; March 2003; 3 pp.; In English; No Copyright;Avail.: CASI: A01, Hardcopy
This work is motivated by the need to accurately predict heat transfer in turbomachinery. For efficient gas turbine operation, flow temperatures in the hot gas path exceed acceptable metal temperatures in many regions of the engine. So that the integrity of the parts can be maintained for an acceptable engine life, the parts must be cooled. Efficient cooling schemes require accurate heat transfer prediction to minimize regions that are overcooled and, even more importantly, to ensure adequate cooling in high-heat-flux regions. Derived from text
Turbine Blades; High Temperature Gases; Cooling; Heat Transfer
20050214849 NASA Glenn Research Center, Cleveland, OH, USA
Low NOx, Lean Direct Wall Injection Combustor Concept Developed
Tacina, Robert R.; Wey, Changlie; Choi, Kyung J.; Research and Technology 2002; March 2003; 2 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
The low-emissions combustor development at the NASA Glenn Research Center is directed toward advanced high-pressure aircraft gas turbine applications. The emphasis of this research is to reduce nitrogen oxides (NOx) at high-power conditions and to maintain carbon monoxide and unburned hydrocarbons at their current low levels at low-power conditions. Low-NOx combustors can be classified into rich burn and lean burn concepts. Lean burn combustors can be further classified into lean-premixed-prevaporized (LPP) and lean direct injection (LDI) combustors. In both concepts, all the combustor air, except for liner cooling flow, enters through the combustor dome so that the combustion occurs at the lowest possible flame temperature. The LPP concept has been shown to have the lowest NOx emissions, but for advanced high-pressure-ratio engines, the possibly of autoignition or flashback precludes its use. LDI differs from LPP in that the fuel is injected directly into the flame zone and, thus, does not have the potential for autoignition or flashback and should have greater stability. However, since it is not premixed and prevaporized, the key is good atomization and mixing of the fuel quickly and uniformly so that flame temperatures are low and NOx formation levels are comparable to those of LPP. Derived from text
Nitrogen Oxides; Emission; Combustion Chambers; Gas Turbines; Compressed Gas
20050214873 NASA Glenn Research Center, Cleveland, OH, USA, Toledo Univ., OH, USA
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Influence of Shock Wave on the Flutter Behavior of Fan Blades Investigated
Srivastava, Rakesh; Bakhle, Milind A.; Stefko, George L.; Research and Technology 2002; March 2003; 2 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
Modern fan designs have blades with forward sweep; a lean, thin cross section; and a wide chord to improve performance and reduce noise. These geometric features coupled with the presence of a shock wave can lead to flutter instability. Flutter is a self-excited dynamic instability arising because of fluid-structure interaction, which causes the energy from the surrounding fluid to be extracted by the vibrating structure. An in-flight occurrence of flutter could be catastrophic and is a significant design issue for rotor blades in gas turbines. Understanding the flutter behavior and the influence of flow features on flutter will lead to a better and safer design. An aeroelastic analysis code, TURBO, has been developed and validated for flutter calculations at the NASA Glenn Research Center. The code has been used to understand the occurrence of flutter in a forward-swept fan design. The forward-swept fan, which consists of 22 inserted blades, encountered flutter during wind tunnel tests at part speed conditions. Derived from text
Shock Waves; Flutter; Fan Blades
20050215120 NASA Glenn Research Center, Cleveland, OH, USA
Parametric Investigation of Thrust Augmentation by Ejectors on a Pulsed Detonation Tube
Wilson, Jack; Sgondea, Alexandru; Paxson, Daniel E.; Rosenthal, Bruce N.; August 2005; 16 pp.; In English; 41st Joint Propulsion Conference, 10-13 Jul. 2005, Tucson, AZ, USA Contract(s)/Grant(s): NAS3-00145; WBS 22-708-03-05 Report No.(s): NASA/TM-2005-213823; AIAA Paper 2005-4208; E-15182; Copyright; Avail.: CASI: A03, Hardcopy
A parametric investigation has been made of thrust augmentation of a 1 inch diameter pulsed detonation tube by ejectors. A set of ejectors was used which permitted variation of the ejector length, diameter, and nose radius, according to a statistical design of experiment scheme. The maximum augmentations for each ejector were fitted using a polynomial response surface, from which the optimum ejector diameters, and nose radius, were found. Thrust augmentations above a factor of 2 were measured. In these tests, the pulsed detonation device was run on approximately stoichiometric air-hydrogen mixtures, at a frequency of 20 Hz. Later measurements at a frequency of 40 Hz gave lower values of thrust augmentation. Measurements of thrust augmentation as a function of ejector entrance to detonation tube exit distance showed two maxima, one with the ejector entrance upstream, and one downstream, of the detonation tube exit. A thrust augmentation of 2.5 was observed using a tapered ejector. Author
Ejectors; Thrust Augmentation; Pulse Detonation Engines; Experiment Design; Pipes (Tubes); Parameterization
20050215166 NASA Glenn Research Center, Cleveland, OH, USA
Supersonic Rocket Thruster Flow Predicted by Numerical Simulation
Davoudzadeh, Farhad; Research and Technology 2003; May 2004; 4 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
Despite efforts in the search for alternative means of energy, combustion still remains the key source. Most propulsion systems primarily use combustion for their needed thrust. Associated with these propulsion systems are the high-velocity hot exhaust gases produced as the byproducts of combustion. These exhaust products often apply uneven high temperature and pressure over the surfaces of the appended structures exposed to them. If the applied pressure and temperature exceed the design criteria of the surfaces of these structures, they will not be able to protect the underlying structures, resulting in the failure of the vehicle mission. An understanding of the flow field associated with hot exhaust jets and the interactions of these jets with the structures in their path is critical not only from the design point of view but for the validation of the materials and manufacturing processes involved in constructing the materials from which the structures in the path of these jets are made. The hot exhaust gases often flow at supersonic speeds, and as a result, various incident and reflected shock features are present. These shock structures induce abrupt changes in the pressure and temperature distribution that need to be considered. In addition, the jet flow creates a gaseous plume that can easily be traced from large distances. To study the flow field associated with the supersonic gases induced by a rocket engine, its interaction with the surrounding surfaces, and its effects on the strength and durability of the materials exposed to it, NASA Glenn Research Center s Combustion Branch teamed with the Ceramics Branch to provide testing and analytical support. The experimental work included the full range of heat flux environments that the rocket engine can produce over a flat specimen. Chamber pressures were varied from 130 to 500 psia and oxidizer-to-fuel ratios (o/f) were varied from 1.3 to 7.5. Author
Supersonic Speed; Thrust Chambers; Numerical Analysis; Combustion Products; Design Analysis
20050215216 NASA Glenn Research Center, Cleveland, OH, USA
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Impulsive Injection for Compressor Stator Separation Control
Culley, Dennis E.; Braunscheidel, Edward P.; Bright, Michelle M.; August 2005; 17 pp.; In English; AIAA Joint Propulsion Conference and Exhibit, 10-13 Jul. 2005, Tucson, AZ, USA Contract(s)/Grant(s): WBS 22-714-70-52 Report No.(s): NASA/TM-2005-213859; AIAA Paper 2005-3633; E-15229; No Copyright; Avail.: CASI: A03, Hardcopy
Flow control using impulsive injection from the suction surface of a stator vane has been applied in a low speed axial compressor. Impulsive injection is shown to significantly reduce separation relative to steady injection for vanes that were induced to separate by an increase in vane stagger angle of 4 degrees. Injected flow was applied to the airfoil suction surface using spanwise slots pitched in the streamwise direction. Injection was limited to the near-hub region, from 10 to 36 percent of span, to affect the dominant loss due to hub leakage flow. Actuation was provided externally using high-speed solenoid valves closely coupled to the vane tip. Variations in injected mass, frequency, and duty cycle are explored. The local corrected total pressure loss across the vane at the lower span region was reduced by over 20 percent. Additionally, low momentum fluid migrating from the hub region toward the tip was effectively suppressed resulting in an overall benefit which reduced corrected area averaged loss through the passage by 4 percent. The injection mass fraction used for impulsive actuation was typically less than 0.1 percent of the compressor through flow. Author
Compressors; Injection; Stators; Vanes; Suction; Airfoils
20050215242 NASA Glenn Research Center, Cleveland, OH, USA
Probability Distribution Estimated From the Minimum, Maximum, and Most Likely Values: Applied to Turbine Inlet Temperature Uncertainty
Holland, Frederic A., Jr.; Research and Technology 2003; May 2004; 2 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
Modern engineering design practices are tending more toward the treatment of design parameters as random variables as opposed to fixed, or deterministic, values. The probabilistic design approach attempts to account for the uncertainty in design parameters by representing them as a distribution of values rather than as a single value. The motivations for this effort include preventing excessive overdesign as well as assessing and assuring reliability, both of which are important for aerospace applications. However, the determination of the probability distribution is a fundamental problem in reliability analysis.
A random variable is often defined by the parameters of the theoretical distribution function that gives the best fit to experimental data. In many cases the distribution must be assumed from very limited information or data. Often the types of information that are available or reasonably estimated are the minimum, maximum, and most likely values of the design parameter. For these situations the beta distribution model is very convenient because the parameters that define the distribution can be easily determined from these three pieces of information. Widely used in the field of operations research, the beta model is very flexible and is also useful for estimating the mean and standard deviation of a random variable given only the aforementioned three values. However, an assumption is required to determine the four parameters of the beta distribution from only these three pieces of information (some of the more common distributions, like the normal, lognormal, gamma, and Weibull distributions, have two or three parameters).
The conventional method assumes that the standard deviation is a certain fraction of the range. The beta parameters are then determined by solving a set of equations simultaneously. A new method developed in-house at the NASA Glenn Research Center assumes a value for one of the beta shape parameters based on an analogy with the normal distribution (ref.1). This new approach allows for a very simple and direct algebraic solution without restricting the standard deviation. The beta parameters obtained by the new method are comparable to the conventional method (and identical when the distribution is symmetrical). However, the proposed method generally produces a less peaked distribution with a slightly larger standard deviation (up to 7 percent) than the conventional method in cases where the distribution is asymmetric or skewed. The beta distribution model has now been implemented into the Fast Probability Integration (FPI) module used in the NESSUS computer code for probabilistic analyses of structures (ref. 2). Author
Normal Density Functions; Probability Theory; Distribution Functions; Inlet Temperature; Turbines; Design Analysis; Standard Deviation; Reliability Analysis
20050215244 NASA Glenn Research Center, Cleveland, OH, USA
Structural Benchmark Testing of Superalloy Lattice Block Subelements Completed Research and Technology
2003; May 2004; 3 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
Superalloy lattice block panels, which are produced directly by investment casting, are composed of thin ligaments arranged in three-dimensional triangulated trusslike structures (see the preceding figure). Optionally, solid panel face sheets can be formed integrally during casting. In either form, lattice block panels can easily be produced with weights less than 25 percent of the mass of a solid panel. Inconel 718 (IN 718) and MarM-247 superalloy lattice block panels have been developed under NASA's Ultra-Efficient Engine Technology Project and Higher Operating Temperature Propulsion Components Project to take advantage of the superalloys' high strength and elevated temperature capability with the inherent light weight and high stiffness of the lattice architecture (ref. 1). These characteristics are important in the future development of turbine engine components. Casting quality and structural efficiency were evaluated experimentally using small beam specimens machined from the cast and heat treated 140- by 300- by 11-mm panels. The matrix of specimens included samples of each superalloy in both open-celled and single-face-sheet configurations, machined from longitudinal, transverse, and diagonal panel orientations. Thirty-five beam subelements were tested in Glenn's Life Prediction Branch's material test machine at room temperature and 650 C under both static (see the following photograph) and cyclic load conditions. Surprisingly, test results exceeded initial linear elastic analytical predictions. This was likely a result of the formation of plastic hinges and redundancies inherent in lattice block geometry, which was not considered in the finite element models. The value of a single face sheet was demonstrated by increased bending moment capacity, where the face sheet simultaneously increased the gross section modulus and braced the compression ligaments against early buckling as seen in open-cell specimens. Preexisting flaws in specimens were not a discriminator in flexural, shear, or stiffness measurements, again because of redundant load paths available in the lattice block structure. Early test results are available in references 2 and 3; more complete analyses are scheduled for publication in 2004. Author
Heat Resistant Alloys; High Temperature; Investment Casting; Gas Turbine Engines; Triangulation; Operating Temperature; Ligaments; High Strength
20050215262 NASA Glenn Research Center, Cleveland, OH, USA
Cryogenic Electric Motor Tested
Brown, Gerald V.; Research and Technology 2003; May 2004; 2 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
Technology for pollution-free ‘electric flight' is being evaluated in a number of NASA Glenn Research Center programs. One approach is to drive propulsive fans or propellers with electric motors powered by fuel cells running on hydrogen. For large transport aircraft, conventional electric motors are far too heavy to be feasible. However, since hydrogen fuel would almost surely be carried as liquid, a propulsive electric motor could be cooled to near liquid hydrogen temperature (-423 F) by using the fuel for cooling before it goes to the fuel cells. Motor windings could be either superconducting or high purity normal copper or aluminum. The electrical resistance of pure metals can drop to 1/100th or less of their room-temperature resistance at liquid hydrogen temperature. In either case, super or normal, much higher current density is possible in motor windings. This leads to more compact motors that are projected to produce 20 hp/lb or more in large sizes, in comparison to on the order of 2 hp/lb for large conventional motors. High power density is the major goal. To support cryogenic motor development, we have designed and built in-house a small motor (7-in. outside diameter) for operation in liquid nitrogen. Author
Electric Motors; Cryogenics; Superconductivity; Cooling; Hydrogen Fuels; Electrical Resistance
20050215271 NASA Glenn Research Center, Cleveland, OH, USA
Reduced-Order Blade Mistuning Analysis Techniques Developed for the Robust Design of Engine Rotors
Min, James B.; Research and Technology 2003; May 2004; 4 pp.; In English; Original contains color and black and white illustrations; No Copyright; Avail.: CASI: A01, Hardcopy
The primary objective of this research program is to develop vibration analysis tools, design tools, and design strategies to significantly improve the safety and robustness of turbine engine rotors. Bladed disks in turbine engines always feature small, random blade-to-blade differences, or mistuning. Mistuning can lead to a dramatic increase in blade forced-response amplitudes and stresses. Ultimately, this results in high-cycle fatigue, which is a major safety and cost concern. In this research program, the necessary steps will be taken to transform a state-of-the-art vibration analysis tool, the Turbo-Reduce forced-response prediction code, into an effective design tool by enhancing and extending the underlying modeling and analysis methods. Furthermore, novel techniques will be developed to assess the safety of a given design. In particular, a procedure will be established for using eigenfrequency curve veerings to identify ‘danger zones' in the operating conditions--ranges of rotational speeds and engine orders in which there is a great risk that the rotor blades will suffer high stresses. This work also will aid statistical studies of the forced response by reducing the necessary number of simulations. Finally, new strategies for improving the design of rotors will be pursued. Several methods will be investigated, including the use of intentional mistuning patterns to mitigate the harmful effects of random mistuning, and the modification of disk stiffness to avoid reaching critical values of interblade coupling in the desired operating range. Recent research progress is summarized in the following paragraphs. First, significant progress was made in the development of the component mode mistuning (CMM) and static mode compensation (SMC) methods for reduced-order modeling of mistuned bladed disks (see the following figure). The CMM method has been formalized and extended to allow a general treatment of mistuning. In addition, CMM allows individual mode mistuning, which accounts for the realistic effects of local variations in blade properties that lead to different mistuning values for different mode types (e.g., mistuning of the first torsion mode versus the second flexural mode). The accuracy and efficiency of the CMM method and the corresponding Turbo-Reduce code were validated for an example finite element model of a bladed disk. Author
Dynamic Structural Analysis; Turbine Engines; Rotors; Robustness (Mathematics); Finite Element Method; Mathematical Models
20050215281 NASA Glenn Research Center, Cleveland, OH, USA, Toledo Univ., OH, USA
NASA Glenn's Seals Group Inaugurated a New State-of-the-Art High-Temperature Test Rig
DeMange, Jeffrey J.; Dunlap, Patrick H., Jr.; Steinetz, Bruce M.; Research and Technology 2003; May 2004; 3 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
The NASA Glenn Research Center is developing advanced control surface seals and propulsion system seals for future space and launch vehicles. To evaluate new seal designs, the Glenn Seals Team recently inaugurated a new state-of-the-art high temperature seal test facility. The Hot Compression/Hot Scrub Rig can perform either high-temperature seal-compression tests or scrub tests at temperatures of up to 3000 F by using different combinations of test fixtures made of monolithic silicon carbide (Hexoloy alpha-SiC), as shown in the following figures. For lower temperature tests (up to 1500 F), Inconel X-750 test fixturing can be used. Derived from text
Compression Tests; High Temperature Tests; Inconel (Trademark); Silicon Carbides
20050215285 NASA Glenn Research Center, Cleveland, OH, USA, Analex Corp., USA, ZIN Technologies, Inc., USA
Testing of Composite Fan Vanes With Erosion-Resistant Coating Accelerated
Bowman, Cheryl L.; Sutter, James K.; Otten, Kim D.; Samorezov, Sergey; Perusek, Gail P.; Research and Technology 2003; May 2004; 3 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
The high-cycle fatigue of composite stator vanes provided an accelerated life-state prior to insertion in a test stand engine. The accelerated testing was performed in the Structural Dynamics Laboratory at the NASA Glenn Research Center under the guidance of Structural Mechanics and Dynamics Branch personnel. Previous research on fixturing and test procedures developed at Glenn determined that engine vibratory conditions could be simulated for polymer matrix composite vanes by using the excitation of a combined slip table and electrodynamic shaker in Glenn's Structural Dynamics Laboratory. Bench-top testing gave researchers the confidence to test the coated vanes in a full-scale engine test. Author
Accelerated Life Tests; Coating; Dynamic Structural Analysis; Erosion; Polymer Matrix Composites
20050215287 NASA Glenn Research Center, Cleveland, OH, USA, Army Research Lab., Cleveland, OH, USA
'Fan-Tip-Drive' High-Power-Density, Permanent Magnet Electric Motor and Test Rig Designed for a Nonpolluting Aircraft Propulsion Program
Brown, Gerald V.; Kascak, Albert F.; Research and Technology 2003; May 2004; 2 pp.; In English; Original contains color illustrations; No Copyright; Avail.: CASI: A01, Hardcopy
A scaled blade-tip-drive test rig was designed at the NASA Glenn Research Center. The rig is a scaled version of a direct-current brushless motor that would be located in the shroud of a thrust fan. This geometry is very attractive since the allowable speed of the armature is approximately the speed of the blade tips (Mach 1 or 1100 ft/s). The magnetic pressure generated in the motor acts over a large area and, thus, produces a large force or torque. This large force multiplied by the large velocity results in a high-power-density motor. Derived from text
Electric Motors; Transonic Speed; Scale Models; Blade Tips
20050215564 NASA Glenn Research Center, Cleveland, OH, USA
Numerical Simulation of the RTA Combustion Rig
Davoudzadeh, Farhad; Buehrle, Robert; Liu, Nan-Suey; Winslow, Ralph; August 2005; 22 pp.; In English; 40th Combustion, 28th Airbreathing Propulsion, 22nd Propulsion Systems Hazards, 4th Modeling and Simulations Joint Subcommittees Meetings, 13-17 Jun. 2005, Charleston, SC, USA Contract(s)/Grant(s): WBS 22-714-02-20 Report No.(s): NASA/TM-2005-213899; E-15270; No Copyright; Avail.: CASI: A03, Hardcopy
The Revolutionary Turbine Accelerator (RTA)/Turbine Based Combined Cycle (TBCC) project is investigating turbine-based propulsion systems for access to space. NASA Glenn Research Center and GE Aircraft Engines (GEAE) planned to develop a ground demonstrator engine for validation testing. The demonstrator (RTA-1) is a variable cycle, turbofan ramjet designed to transition from an augmented turbofan to a ramjet that produces the thrust required to accelerate the vehicle from Sea Level Static (SLS) to Mach 4. The RTA-1 is designed to accommodate a large variation in bypass ratios from sea level static to Mach 4 conditions. Key components of this engine are new, such as a nickel alloy fan, advanced trapped vortex combustor, a Variable Area Bypass Injector (VABI), radial flameholders, and multiple fueling zones. A means to mitigate risks to the RTA development program was the use of extensive component rig tests and computational fluid dynamics (CFD) analysis. Author
Trapped Vortices; Flame Holders; Computational Fluid Dynamics; Combustion; Turbines; Propulsion System Configurations; Injectors
20050215595 NASA Glenn Research Center, Cleveland, OH, USA
Magnetically Levitated Ducted Fan Being Developed as a Propulsor Option for Electric Flight
Emerson, Dawn C.; Research and Technology 2003; May 2004; 3 pp.; In English; Original contains color illustrations; No Copyright; Avail.: CASI: A01, Hardcopy
The NASAGlenn Research Center is developing a revolutionary engine concept called the Magnetically Levitated Ducted Fan. The objective of this work is to demonstrate the feasibility of a pollution-free electromagnetic propulsor for all-electric flight. This concept will help to reduce harmful emissions, to reduce the Nation's dependence on fossil fuels, and to mitigate many of the concerns and limitations encountered in conventional engine systems such as bearing wear, leaks, seal failure, and friction loss. Author
Engine Design; Ducted Fans; Magnetic Suspension; Friction
20050215597 NASA Glenn Research Center, Cleveland, OH, USA
Advanced Manufacturing Techniques Demonstrated for Fabricating Developmental Hardware
Redding, Chip; Research and Technology 2003; May 2004; 2 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
NASA Glenn Research Center's Engineering Development Division has been working in support of innovative gas turbine engine systems under development by Glenn's Combustion Branch. These one-of-a-kind components require operation under extreme conditions. High-temperature ceramics were chosen for fabrication was because of the hostile operating environment. During the designing process, it became apparent that traditional machining techniques would not be adequate to produce the small, intricate features for the conceptual design, which was to be produced by stacking over a dozen thin layers with many small features that would then be aligned and bonded together into a one-piece unit. Instead of using traditional machining, we produced computer models in Pro/ENGINEER (Parametric Technology Corporation (PTC), Needham, MA) to the specifications of the research engineer. The computer models were exported in stereolithography standard (STL) format and used to produce full-size rapid prototype polymer models. These semi-opaque plastic models were used for visualization and design verification. The computer models also were exported in International Graphics Exchange Specification (IGES) format and sent to Glenn's Thermal/Fluids Design &Analysis Branch and Applied Structural Mechanics Branch for profiling heat transfer and mechanical strength analysis. Author
Design Analysis; Gas Turbine Engines; Systems Engineering; Manufacturing; Product Development; High Temperature; Engine Design; Mechanics (Physics)
20050215640 NASA Glenn Research Center, Cleveland, OH, USA
Photochemically Etched Construction Technology Developed for Digital Xenon Feed Systems
Otsap, Ben; Cardin, Joseph; Verhey, Timothy R.; Rawlin, Vincent K.; Mueller, Juergen; Aadlund, Randall; Kay, Robert; Andrews, Michael; Research and Technology 2003; May 2004; 4 pp.; In English; No Copyright;Avail.: CASI: A01, Hardcopy
Electric propulsion systems are quickly emerging as attractive options for primary propulsion in low Earth orbit, in geosynchronous orbit, and on interplanetary spacecraft. The driving force behind the acceptance of these systems is the substantial reduction in the propellant mass that can be realized. Unfortunately, system designers are often forced to utilize components designed for chemical propellants in their electric systems. Although functionally acceptable, these relatively large, heavy components are designed for the higher pressures and mass flow rates required by chemical systems. To fully realize the benefits of electric propulsion, researchers must develop components that are optimized for the low flow rates, critical leakage needs, low pressures, and limited budgets of these emerging systems. Derived from text
Technology Assessment; Photochemical Reactions; Xenon; Feed Systems; Propulsion System Performance; Electric Propulsion
20050215686 NASA Glenn Research Center, Cleveland, OH, USA
Low-Speed Active Flow Control Laboratory Developed Culley, Dennis E.; Bright, Michelle M.; Research and Technology 2004; June 2005; 3 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
The future of aviation propulsion systems is increasingly focused on the application of control technologies to significantly enhance the performance of a new generation of air vehicles. Active flow control refers to a set of technologies that manipulate the flow of air and combustion gases deep within the confines of an engine to dynamically alter its performance during flight. By employing active flow control, designers can create engines that are significantly lighter, are more fuel efficient, and produce lower emissions. In addition, the operating range of an engine can be extended, yielding safer transportation systems. The realization of these future propulsion systems requires the collaborative development of many base technologies to achieve intelligent, embedded control at the engine locations where it will be most effective. NASA Glenn Research Center s Controls and Dynamics Technology Branch has developed a state-of-the-art low-speed Active Flow Control Laboratory in which emerging technologies can be integrated and explored in a flexible, low-cost environment. The facility allows the most promising developments to be prescreened and optimized before being tested on higher fidelity platforms, thereby reducing the cost of experimentation and improving research effectiveness. Derived from text
Propulsion System Configurations; Propulsion System Performance; Active Control; Air Flow; Low Speed
20050215690 NASA Glenn Research Center, Cleveland, OH, USA
Investigation of Exoskeletal Engine Propulsion System Concept
Roche, Joseph M.; Palac, Donald T.; Hunter, James E.; Myers, David E.; Snyder, Christopher A.; Kosareo, Daniel N.; McCurdy, David R.; Dougherty, Kevin T.; August 2005; 101 pp.; In English; Original contains color and black and white illustrations Contract(s)/Grant(s): WBS 22-708-87-07 Report No.(s): NASA/TM-2005-213369; E-14837; No Copyright; Avail.: CASI: A06, Hardcopy
An innovative approach to gas turbine design involves mounting compressor and turbine blades to an outer rotating shell. Designated the exoskeletal engine, compression (preferable to tension for high-temperature ceramic materials, generally) becomes the dominant blade force. Exoskeletal engine feasibility lies in the structural and mechanical design (as opposed to cycle or aerothermodynamic design), so this study focused on the development and assessment of a structural-mechanical exoskeletal concept using the Rolls-Royce AE3007 regional airliner all-axial turbofan as a baseline. The effort was further limited to the definition of an exoskeletal high-pressure spool concept, where the major structural and thermal challenges are represented. The mass of the high-pressure spool was calculated and compared with the mass of AE3007 engine components. It was found that the exoskeletal engine rotating components can be significantly lighter than the rotating components of a conventional engine. However, bearing technology development is required, since the mass of existing bearing systems would exceed rotating machinery mass savings. It is recommended that once bearing technology is sufficiently advanced, a ‘clean sheet' preliminary design of an exoskeletal system be accomplished to better quantify the potential for the exoskeletal concept to deliver benefits in mass, structural efficiency, and cycle design flexibility. Author
Propulsion System Configurations; Turbofans; Gas Turbine Engines; Turbine Blades; Compressor Blades; Aerothermodynamics
20050215691 NASA Glenn Research Center, Cleveland, OH, USA
Intelligent Life-Extending Controls for Aircraft Engines Studied
Guo, Ten-Huei; Research and Technology 2004; June 2005; 4 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
Current aircraft engine controllers are designed and operated to provide desired performance and stability margins. Except for the hard limits for extreme conditions, engine controllers do not usually take engine component life into consideration during the controller design and operation. The end result is that aircraft pilots regularly operate engines under unnecessarily harsh conditions to strive for optimum performance. The NASA Glenn Research Center and its industrial and academic partners have been working together toward an intelligent control concept that will include engine life as part of the controller design criteria. This research includes the study of the relationship between control action and engine component life as well as the design of an intelligent control algorithm to provide proper tradeoffs between performance and engine life. This approach is expected to maintain operating safety while minimizing overall operating costs. In this study, the thermomechanical fatigue (TMF) of a critical component was selected to demonstrate how an intelligent engine control algorithm can significantly extend engine life with only a very small sacrifice in performance. An intelligent engine control scheme based on modifying the high-pressure spool speed (NH) was proposed to reduce TMF damage from ground idle to takeoff. The NH acceleration schedule was optimized to minimize the TMF damage for a given rise-time constraint, which represents the performance requirement. The intelligent engine control scheme was used to simulate a commercial short-haul aircraft engine. Derived from text
Algorithms; Controllers; Damage; Engine Control; Thermodynamics
20050215694 NASA Glenn Research Center, Cleveland, OH, USA, Army Research Lab., Cleveland, OH, USA
Discrete Event Supervisory Control Applied to Propulsion Systems
Litt, Jonathan S.; Shah, Neerav; Research and Technology 2004; June 2005; 3 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
The theory of discrete event supervisory (DES) control was applied to the optimal control of a twin-engine aircraft propulsion system and demonstrated in a simulation. The supervisory control, which is implemented as a finite-state automaton, oversees the behavior of a system and manages it in such a way that it maximizes a performance criterion, similar to a traditional optimal control problem. DES controllers can be nested such that a high-level controller supervises multiple lower level controllers. This structure can be expanded to control huge, complex systems, providing optimal performance and increasing autonomy with each additional level. The DES control strategy for propulsion systems was validated using a distributed testbed consisting of multiple computers--each representing a module of the overall propulsion system--to simulate real-time hardware-in-the-loop testing. In the first experiment, DES control was applied to the operation of a nonlinear simulation of a turbofan engine (running in closed loop using its own feedback controller) to minimize engine structural damage caused by a combination of thermal and structural loads. This enables increased on-wing time for the engine through better management of the engine-component life usage. Thus, the engine-level DES acts as a life-extending controller through its interaction with and manipulation of the engine s operation. Derived from text
Complex Systems; Controllers; Engine Parts; Feedback Control; Propulsion System Performance; Turbofan Engines
20050216392 NASA Glenn Research Center, Cleveland, OH, USA
Ejector Enhanced Pulsejet Based Pressure Gain Combustors: An Old Idea With a New Twist
Paxson, Daniel E.; Dougherty, Kevin T.; August 2005; 22 pp.; In English; 41st Joint Propulsion Conference and Exhibit, 10-13 Jul. 2005, Tucson, AZ, USA; Original contains color and black and white illustrations Contract(s)/Grant(s): WBS 22-066-10-05 Report No.(s): NASA/TM-2005-213854; E-15224; AIAA Paper 2005-4216; No Copyright; Avail.: CASI: A03, Hardcopy
An experimental investigation of pressure-gain combustion for gas turbine application is described. The test article consists of an off-the-shelf valved pulsejet, and an optimized ejector, both housed within a shroud. The combination forms an effective can combustor across which there is a modest total pressure rise rather than the usual loss found in conventional combustors. Although the concept of using a pulsejet to affect semi-constant volume (i.e., pressure-gain) combustion is not new, that of combining it with a well designed ejector to efficiently mix the bypass flow is. The result is a device which to date has demonstrated an overall pressure rise of approximately 3.5 percent at an overall temperature ratio commensurate with modern gas turbines. This pressure ratio is substantially higher than what has been previously reported in pulsejet-based combustion experiments. Flow non-uniformities in the downstream portion of the device are also shown to be substantially reduced compared to those within the pulsejet itself. The standard deviation of total pressure fluctuations, measured just downstream of the ejector was only 5.0 percent of the mean. This smoothing aspect of the device is critical to turbomachinery applications since turbine performance is, in general, negatively affected by flow non-uniformities and unsteadiness. The experimental rig will be described and details of the performance measurements will be presented. Analyses showing the thermodynamic benefits from this level of pressure-gain performance in a gas turbine will also be assessed for several engine types. Issues regarding practical development of such a device are discussed, as are potential emissions reductions resulting from the rich burning nature of the pulsejet and the rapid mixing (quenching) associated with unsteady ejectors. Author
Combustion Chambers; Ejectors; Pressure Oscillations; Pulsejet Engines; Pulse Detonation Engines
20050216394 NASA Glenn Research Center, Cleveland, OH, USA
Control Algorithms and Simulated Environment Developed and Tested for Multiagent Robotics for Autonomous Inspection of Propulsion Systems
Wong, Edmond; Research and Technology 2004; June 2005; 4 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
The NASA Glenn Research Center and academic partners are developing advanced multiagent robotic control algorithms that will enable the autonomous inspection and repair of future propulsion systems. In this application, on-wing engine inspections will be performed autonomously by large groups of cooperative miniature robots that will traverse the surfaces of engine components to search for damage. The eventual goal is to replace manual engine inspections that require expensive and time-consuming full engine teardowns and allow the early detection of problems that would otherwise result in catastrophic component failures. As a preliminary step toward the long-term realization of a practical working system, researchers are developing the technology to implement a proof-of-concept testbed demonstration. In a multiagent system, the individual agents are generally programmed with relatively simple controllers that define a limited set of behaviors. However, these behaviors are designed in such a way that, through the localized interaction among individual agents and between the agents and the environment, they result in self-organized, emergent group behavior that can solve a given complex problem, such as cooperative inspection. One advantage to the multiagent approach is that it allows for robustness and fault tolerance through redundancy in task handling. In addition, the relatively simple agent controllers demand minimal computational capability, which in turn allows for greater miniaturization of the robotic agents. Derived from text
Algorithms; Computer Systems Performance; Autonomy; Controllers; Propulsion System Performance
20050216399 NASA Glenn Research Center, Cleveland, OH, USA
Planar Inlet Design and Analysis Process (PINDAP)
Slater, JohnW.; Gruber, Christopher R.; August 2005; 17 pp.; In English; 41st Joint Propulsion Conference and Exhibit, 10-13 Jul. 2005, Tucson,AZ, USA; Original contains color and black and white illustrations Contract(s)/Grant(s): WBS 22-714-09-25 Report No.(s): NASA/TM-2005-213866; E-15236; AIAA Paper 2005-4203; No Copyright; Avail.: CASI: A03, Hardcopy
The Planar Inlet Design and Analysis Process (PINDAP) is a collection of software tools that allow the efficient aerodynamic design and analysis of planar (two-dimensional and axisymmetric) inlets. The aerodynamic analysis is performed using the Wind-US computational fluid dynamics (CFD) program. A major element in PINDAP is a Fortran 90 code named PINDAP that can establish the parametric design of the inlet and efficiently model the geometry and generate the grid for CFD analysis with design changes to those parameters. The use of PINDAP is demonstrated for subsonic, supersonic, and hypersonic inlets. Author
Computational Fluid Dynamics; Hypersonic Inlets; Supersonic Inlets; Planar Structures; Engine Design
20050217135 Japan Aerospace Exploration Agency, Japan
Overview of Japanese Research and Development Program on Liquid Rocket Engine Combustor
Hasegawa, K.; Kumakawa, A.; Onodera, T.; Shimizu, T.;Watanabe, Y.; Tomita, T.; Taniguchi, H.; Naruo,Y.; Fifth International Symposium on Liquid Space Propulsion; January 2005; 39 pp.; In English; See also 20050217123; Original contains color illustrations; No Copyright; Avail.: CASI: A03, Hardcopy
Contents include the following: (1) Laser Ablation Ignition. (2) Extendible Nozzle and Dual-Bell Nozzle. (3) LE-7A Nozzle Flow Separation Phenomenon. (4) Low Cycle Chamber Pressure Fluctuation. (5) LOX/LH2 Engine for Reusable Vehicle Test. (6) Study of Advanced Expander Bleed Cycle. CASI
Combustion Chambers; Engine Tests; Ignition; Hydrogen Oxygen Engines; Nozzle Flow
Source: NASA.
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