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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 1,423 records · Page 79
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NASA NTRS2001

New Technology Sparks Smoother Engines and Cleaner Air

Automotive Resources, Inc. (ARI) has developed a new device for igniting fuel in engines-the SmartPlug.TM SmartPlug is a self-contained ignition system that may be retrofitted to existing spark-ignition and compression-ignition engines. The SmartPlug needs as little as six watts of power for warm-up, and requires no electricity at all when the engine is running. Unlike traditional spark plugs, once the SmartPlug ignites the engine, and the engine heats up, the power supply for the plug is no longer necessary. In the utility industry, SmartPlugs can be used in tractors, portable generators, compressors, and pumps. In addition to general-purpose applications, such as lawn mowers and chainsaws, SmartPlugs can also be used in the recreational, marine, aviation, and automotive industries. Unlike traditional ignition systems, the SmartPlug system requires no distributor, coil points, or moving parts. SmartPlugs are non-fouling, with a faster and cleaner burn than traditional spark plugs. They prevent detonation and are not sensitive to moisture, allowing them to be used on a variety of engines. Other advantages include no electrical noise, no high voltage, exceptionally high altitude capabilities, and better cold-start statistics than those of standard spark ignition systems. Future applications for the SmartPlug are being evaluated by manufacturers in the snowmobile industry.

Source record↗
NASA NTRS2002

Ultra Efficient Engine Technology Systems Integration and Environmental Assessment

This study documents the design and analysis of four types of advanced technology commercial transport airplane configurations (small, medium large and very large) with an assumed technology readiness date of 2010. These airplane configurations were used as a platform to evaluate the design concept and installed performance of advanced technology engines being developed under the NASA Ultra Efficient Engine Technology (UEET) program. Upon installation of the UEET engines onto the UEET advanced technology airframes, the small and medium airplanes both achieved an additional 16% increase in fuel efficiency when using GE advanced turbofan engines. The large airplane achieved an 18% increase in fuel efficiency when using the P&W geared fan engine. The very large airplane (i.e. BWB), also using P&W geared fan engines, only achieved an additional 16% that was attributed to a non-optimized airplane/engine combination.

Daggett, David L.↗
NASA NTRS2002

ISO 9000 and/or Systems Engineering Capability Maturity Model?

For businesses and organizations to remain competitive today they must have processes and systems in place that will allow them to first identify customer needs and then develop products/processes that will meet or exceed the customers needs and expectations. Customer needs, once identified, are normally stated as requirements. Designers can then develop products/processes that will meet these requirements. Several functions, such as quality management and systems engineering management are used to assist product development teams in the development process. Both functions exist in all organizations and both have a similar objective, which is to ensure that developed processes will meet customer requirements. Are efforts in these organizations being duplicated? Are both functions needed by organizations? What are the similarities and differences between the functions listed above? ISO 9000 is an international standard of goods and services. It sets broad requirements for the assurance of quality and for management's involvement. It requires organizations to document the processes and to follow these documented processes. ISO 9000 gives customers assurance that the suppliers have control of the process for product development. Systems engineering can broadly be defined as a discipline that seeks to ensure that all requirements for a system are satisfied throughout the life of the system by preserving their interrelationship. The key activities of systems engineering include requirements analysis, functional analysis/allocation, design synthesis and verification, and system analysis and control. The systems engineering process, when followed properly, will lead to higher quality products, lower cost products, and shorter development cycles. The System Engineering Capability Maturity Model (SE-CMM) will allow companies to measure their system engineering capability and continuously improve those capabilities. ISO 9000 and SE-CMM seem to have a similar objective, which is to document the organization's processes and certify to potential customers the capability of a supplier to control the processes that determine the quality of the product or services being produced. The remaining sections of this report examine the differences and similarities between ISO 9000 and SE-CMM and make recommendations for implementation.

Sampson E. Gholston↗
NASA NTRS2002

Extending Ion Engine Technology to NEXT and Beyond

Extending ion engine technology beyond the current state-of-the art primary interplanetary electric propulsion system, the 2.3-kW NASA Solar Electric Propulsion Technology and Applications Readiness (NSTAR) system, will require thrusters with improved propellant throughput and total impulse capability. Many of the design choices that culminated in the NSTAR thrusters must be revisited, and their application to next generation ion engine technology must be evaluated. The concept of derating, which was successfully employed in NSTAR, has been applied to the 40 cm NASA Evolutionary Xenon Thruster (NEXT) currently under development at NASA Glenn Research Center (GRC). At 5-kW, NEXT operates with the same average beam current density as NSTAR, and at 10-kW, the peak beam current density is only ten percent greater than NSTAR. The result is that similar Ion optics technology is expected to yield comparable lifetime. Thick-accelerator- grid ion optics are also being tested to realize additional lifetime benefits. A 40-A discharge cathode is being developed for NEXT based on scaling the NSTAR design. Nevertheless, the experiences of the NSTAR ground tests and the thruster on the Deep Space One spacecraft indicate that the discharge cathode wear must be studied experimentally and theoretically to ensure that it meets the lifetime requirements. Although NEXT is in its infancy, investigations have already begun to examine possible modifications to engine design for even higher-power and higher-specific impulse engines. Ion optics using alternate materials such as titanium, graphite, or carbon-carbon composite are currently being investigated due to their low sputter yields at high voltage. To avoid the difficulties encountered using electrodes at high-currents, the use of a microwave-based ion thruster is under investigation for potential high-power ion thruster systems requiring long lifetimes. Additionally, alternative propellants are being considered for applications requiring high-specific impulse (>> 5000 s) and extremely long-life (>> 15,000 hr). Testing requirements make condensable propellants attractive for high-power engines. Although the NSTAR ion engine demonstrated the flight maturity of ion thruster technology, many challenges remain for the development of thrusters with improved propellant throughput and power handling capabilities.

Domonkos, Matthew T.↗
NASA NTRS2002

The Use of Probabilistic Methods to Evaluate the Systems Impact of Component Design Improvements on Large Turbofan Engines

Probabilistic Structural Analysis (PSA) is now commonly used for predicting the distribution of time/cycles to failure of turbine blades and other engine components. These distributions are typically based on fatigue/fracture and creep failure modes of these components. Additionally, reliability analysis is used for taking test data related to particular failure modes and calculating failure rate distributions of electronic and electromechanical components. How can these individual failure time distributions of structural, electronic and electromechanical component failure modes be effectively combined into a top level model for overall system evaluation of component upgrades, changes in maintenance intervals, or line replaceable unit (LRU) redesign? This paper shows an example of how various probabilistic failure predictions for turbine engine components can be evaluated and combined to show their effect on overall engine performance. A generic model of a turbofan engine was modeled using various Probabilistic Risk Assessment (PRA) tools (Quantitative Risk Assessment Software (QRAS) etc.). Hypothetical PSA results for a number of structural components along with mitigation factors that would restrict the failure mode from propagating to a Loss of Mission (LOM) failure were used in the models. The output of this program includes an overall failure distribution for LOM of the system. The rank and contribution to the overall Mission Success (MS) is also given for each failure mode and each subsystem. This application methodology demonstrates the effectiveness of PRA for assessing the performance of large turbine engines. Additionally, the effects of system changes and upgrades, the application of different maintenance intervals, inclusion of new sensor detection of faults and other upgrades were evaluated in determining overall turbine engine reliability.

Packard, Michael H.↗
NASA NTRS2002

Turbine Engine Clearance Control Systems: Current Practices and Future Directions

Improved blade tip sealing in the high pressure compressor (HPC) and high pressure turbine (HPT) can provide dramatic reductions in specific fuel consumption (SFC), time-on-wing, compressor stall margin, and engine efficiency as well as increased payload and mission range capabilities. Maintenance costs to overhaul large commercial gas turbine engines can easily exceed $1M. Engine removal from service is primarily due to spent exhaust gas temperature (EGT) margin caused mainly by the deterioration of HPT components. Increased blade tip clearance is a major factor in hot section component degradation. As engine designs continue to push the performance envelope with fewer parts and the market drives manufacturers to increase service life, the need for advanced sealing continues to grow. A review of aero gas turbine engine HPT performance degradation and the mechanisms that promote these losses are discussed. Benefits to the HPT due to improved clearance management are identified. Past and present sealing technologies are presented along with specifications for next generation engine clearance control systems.

Lattime, Scott B.↗
NASA NTRS2003

Tracking and Control of Gas Turbine Engine Component Damage/Life

This paper describes damage mechanisms and the methods of controlling damages to extend the on-wing life of critical gas turbine engine components. Particularly, two types of damage mechanisms are discussed: creep/rupture and thermo-mechanical fatigue. To control these damages and extend the life of engine hot-section components, we have investigated two methodologies to be implemented as additional control logic for the on-board electronic control unit. This new logic, the life-extending control (LEC), interacts with the engine control and monitoring unit and modifies the fuel flow to reduce component damages in a flight mission. The LEC methodologies were demonstrated in a real-time, hardware-in-the-loop simulation. The results show that LEC is not only a new paradigm for engine control design, but also a promising technology for extending the service life of engine components, hence reducing the life cycle cost of the engine.

Jaw, Link C.↗
NASA NTRS2003

Recent Progress in Engine Noise Reduction Technologies

Highlights from NASA-funded research over the past ten years for aircraft engine noise reduction are presented showing overall technical plans, accomplishments, and selected applications to turbofan engines. The work was sponsored by NASA's Advanced Subsonic Technology (AST) Noise Reduction Program. Emphasis is given to only the engine noise reduction research and significant accomplishments that were investigated at Technology Readiness Levels ranging from 4 to 6. The Engine Noise Reduction sub-element was divided into four work areas: source noise prediction, model scale tests, engine validation, and active noise control. Highlights from each area include technologies for higher bypass ratio turbofans, scarf inlets, forward-swept fans, swept and leaned stators, chevron/tabbed nozzles, advanced noise prediction analyses, and active noise control for fans. Finally, an industry perspective is given from General Electric Aircraft Engines showing how these technologies are being applied to commercial products. This publication contains only presentation vu-graphs from an invited lecture given at the 41st AIAA Aerospace Sciences Meeting, January 6-9, 2003.

Huff, Dennis↗
NASA NTRS1946

Effect of Fuel Composition, Engine Operating Variables, and Spark-Plug Type and Condition on Preignition-Limited Performance of an R-2800 Cylinder

The preignition characteristics of the R-2800 cylinder, as effected by fuel consumption, engine operating variables, and spark plug type and condition, were evaluated. The effects on preignition-limited performance of various percentages of aromatics (benzene, toluene, cumene, xylene) in a base fuel of triptane were investigated. Two paraffins (triptane and S + 6.0 ml TEL/gal) and two refinery blends (28-R and 33-R) were preignition rated. The effect of changes in the following engine operating variables on preignition limit was determined: inlet-air temperature, rear spark plug gasket temperature, engine speed, spark advance, tappet clearance, and oil consumption. Preignition limits of the R-2800 cylinder using Champion C34S and C35S and AC-LS86, LS87, and LS88 spark plugs were established and the effect of spark plug deterioration was investigated. No definite trends in preignition-limited indicated mean effective pressure were indicated for aromatics as a class when increased percentages of different aromatics were added to a base fuel of triptane. Three types of fuel (aromatics, paraffins, and refinery blends) showed a preignition range for this cylinder from 65 to 104 percent when based on the performance of S plus 6.0 ml TEL per gallon as 100 percent. The R-2800 cylinder is therefore relatively insensitive to fuel composition when compared to a CFR F-4 engine, which had a pre-ignition range from 72 to 100 percent for the same fuels. Six engine operating variables were investigated with the following results: preignition-limited indicated mean effective pressure decreased, with increases in engine speed, rear spark plug gasket temperature, inlet-air temperature, and spark advance beyond 20 F B.T.C. and was unaffected by rate of oil consumption or by tappet clearance. Spark plugs were rated over a range of preignition-limited indicated mean effective pressure from 200 to 390 pounds per square inch at a fuel-air ratio of 0.07 in the following order of increased resistance to preignition: AC-LS97, AC-LS88, Champion C358, AC-LS86, and Champion C34S. Spark plug deterioration in the form of cracks in the porcelain had been broken away from the center electrode and were retained in the spark plug cavity, the preignition limit was decreased as much as 57 percent. When the broken pieces had been removed, the preignition limit increased from that of the undamaged porcelain as the weight of removed porcelain was increased.

Pfender, John F.↗
NASA NTRS1948

Preliminary Results of an Altitude-Wind-Tunnel Investigation of an Axial-Flow Gas Turbine-Propeller Engine: Combustion-Chamber Characterisitcs - 5

An investigation to determine the performance and operational characteristics of an axial-flow gas turbine-propeller engine was conducted in the Cleveland altitude wind tunnel. As part of this investigation, the combustion-chamber performance was determined at pressure altitudes from 5000 to 35,000 feet, compressor-inlet ram-pressure ratios of 1.00 and 1.09, and engine speeds from 8000 to 13,000 rpm. Combustion-chamber performance is presented as a function of corrected engine speed and corrected horsepower. For the range of corrected engine speeds investigated, overall total-pressure-loss ratio, cycle efficiency, and the fractional loss in cycle efficiency resulting from pressure losses in the combustion chambers were unaffected by a change in altitude or compressor-inlet ram-pressure ratio. For the range of corrected horsepowers investigated, the total-pressure-loss ratio and the fractional loss in cycle efficiency resulting from pressure losses in the combustion chambers decreased with an increase in corrected horsepower at a constant corrected engine speed. The combustion efficiency remained constant for the range of corrected horsepowers investigated at all corrected engine speeds.

Geisenheyner, Robert M.↗
NASA NTRS2003

COBRA System Engineering Processes to Achieve SLI Strategic Goals

The COBRA Prototype Main Engine Development Project was an endeavor conducted as a joint venture between Pratt & Whitney and Aerojet to conduct risk reduction in LOX/LH2 main engine technology for the NASA Space Launch Initiative (SLI). During the seventeen months of the project (April 2001 to September 2002), approximately seventy reviews were conducted, beginning with the Engine Systems Requirements Review (SRR) and ending with the Engine Systems Interim Design Review (IDR). This paper discusses some of the system engineering practices used to support the reviews and the overall engine development effort.

Ballard, Richard O.↗
NASA NTRS2003

MC-1 Engine Valves, Lessons Learned

Many lessons were learned during the development of the valves for the MC-1 engine. The purpose of this report is to focus on a variety of issues related to the engine valves and convey the lessons learned. This paper will not delve into detailed technical analysis of the components. None of the lessons learned are new or surprising, but simply reinforce the importance of addressing the details of the design early, at the component level. The Marshall Space Flight Center (MSFC), Huntsville, Alabama developed the MC-1 engine, a LOX / FW-1, 60,000 pound thrust engine. This engine was developed under the Low Cost Boost Technology office at MSFC and proved to be a very successful project for the MSFC Propulsion team and the various subcontractors working the development of the engine and its components.

Laszar, John↗
NASA NTRS2003

Towards an Automated Full-Turbofan Engine Numerical Simulation

The objective of this study was to demonstrate the high-fidelity numerical simulation of a modern high-bypass turbofan engine. The simulation utilizes the Numerical Propulsion System Simulation (NPSS) thermodynamic cycle modeling system coupled to a high-fidelity full-engine model represented by a set of coupled three-dimensional computational fluid dynamic (CFD) component models. Boundary conditions from the balanced, steady-state cycle model are used to define component boundary conditions in the full-engine model. Operating characteristics of the three-dimensional component models are integrated into the cycle model via partial performance maps generated automatically from the CFD flow solutions using one-dimensional meanline turbomachinery programs. This paper reports on the progress made towards the full-engine simulation of the GE90-94B engine, highlighting the generation of the high-pressure compressor partial performance map. The ongoing work will provide a system to evaluate the steady and unsteady aerodynamic and mechanical interactions between engine components at design and off-design operating conditions.

Reed, John A.↗
NASA NTRS2003

Independent Review of the Failure Modes of F-1 Engine and Propellants System

The F-1 is the powerful engine, that hurdled the Saturn V launch vehicle from the Earth to the moon on July 16,1969. The force that lifted the rocket overcoming the gravitational force during the first stage of the flight was provided by a cluster of five F-1 rocket engines, each of them developing over 1.5 million pounds of thrust (MSFC-MAN-507). The F-1 Rocket engine used RP-1 (Rocket Propellant-1, commercially known as Kerosene), as fuel with lox (liquid Oxygen) as oxidizer. NASA terminated Saturn V activity and has focused on Space Shuttle since 1972. The interest in rocket system has been revived to meet the National Launch System (NLS) program and a directive from the President to return to the Moon and exploration of the space including Mars. The new program Space Launch Initiative (SLI) is directed to drastically reduce the cost of flight for payloads, and adopt a reusable launch vehicle (RLV). To achieve this goal it is essential to have the ability of lifting huge payloads into low earth orbit. Probably requiring powerful boosters as strap-ons to a core vehicle, as was done for the Saturn launch vehicle. The logic in favor of adopting Saturn system, a proven technology, to meet the SLI challenge is very strong. The F-1 engine was the largest and most powerful liquid rocket engine ever built, and had exceptional performance. This study reviews the failure modes of the F-1 engine and propellant system.

Ray, Paul↗
NASA NTRS2003

Performance Evaluation of the NEXT Ion Engine

The performance test results of three NEXT ion engines are presented. These ion engines exhibited peak specific impulse and thrust efficiency ranges of 4060 4090 s and 0.68 0.69, respectively, at the full power point of the NEXT throttle table. The performance of the ion engines satisfied all project requirements. Beam flatness parameters were significantly improved over the NSTAR ion engine, which is expected to improve accelerator grid service life. The results of engine inlet pressure and temperature measurements are also presented. Maximum main plenum, cathode, and neutralizer pressures were 12,000 Pa, 3110 Pa, and 8540 Pa, respectively, at the full power point of the NEXT throttle table. Main plenum and cathode inlet pressures required about 6 hours to increase to steady-state, while the neutralizer required only about 0.5 hour. Steady-state engine operating temperature ranges throughout the power throttling range examined were 179 303 C for the discharge chamber magnet rings and 132 213 C for the ion optics mounting ring.

Soulas, George C.↗
NASA NTRS2003

Space Shuttle Main Engine Implications for the Abort-to-Orbit Off-the-Pad Study

In 2001, the Space Shuttle Main Engine (SSME) project office was contacted by the Space Shuttle Ascent Guidance, Navigation, and Control group to provide the engine perspective for an Abort-to-Orbit (ATO) study. The purpose of the AT0 Off-the-Pad study was to determine the feasibility of eliminating the Return to Launch Site and Transatlantic abort modes by using a five-segment solid rocket booster and throttling the remaining SSMEs to a higher power level. This would enable all abort modes to be Abort-to-Orbit. The SSME project office at Marshall Space Flight Center collaborated with MSFC's Space Transportation Directorate and Rocketdyne Propulsion and Power to provide the AT0 Off-the-Pad study with the analysis required. Power levels at 109%, 1 11% and 1 13% of rated power level were studied as well as mixture ratio decreases down to 5.85. SSME was to evaluate and define the technical and programmatic impacts to certify the SSME to these abort power levels. The SSME systems analysis group performed a steady state analysis using the SSME power balance model to determine if there were any technical issues associated with higher power level, low mixture ratio operation. Based on each power level/mixture ratio combination, an engine certification plan was created and a preliminary probabilistic risk assessment was performed. The results showed favorable results for higher power 1evel/lower mixture ratio SSME operation. In nearly all performance and redline parameters, the traded engine operation was encompassed by nominal engine performance of a prior engine configuration.

Schoffstoll, Dayna L.↗
NASA NTRS1946

Investigation of Operating Characteristics of an Engine Equipped with Modifications to Eliminate Fuel-Evaporation Icing

Two modified fuel-injection systems, a drilled-inducer type and a spinner type, that prevent serious fuel-evaporation icing were installed on a V-type, liquid-cooled aircraft engine and a preliminary investigation was conducted to determine the effect on engine operating characteristics. The spinner system was also ground - and flight tested on a twin-engine fighter airplane. Flight measurements of cylinder-head temperature over a range of fuel-air ratios and engine power conditions were made at an altitude of approximately 10,000 feet. Starting and accelerating of the engine on the ground were unaffected by the fuel-injection modifications. During the flight investigation, no appreciable variation occurred between the maximum and minimum cylinder-head temperature with the standard and modified system for the same power condition and no irregularity of mixture distribution could be detected throughout the power range of the engine. Normal mixture distribution was also indicated by a similar response of cylinder-head temperature for variations of fuel-air ratio at manifold pressures of 25 and 35 inches of mercury absolute. Both modified fuel-injection systems required less fuel-nozzle pressure than the standard system to obtain the desired fuel-air ratio for given air-flow condition.

Mulholland, Donald R.↗
NASA NTRS2004

Turnaround Time Modeling for Conceptual Rocket Engines

Recent years have brought about a paradigm shift within NASA and the Space Launch Community regarding the performance of conceptual design. Reliability, maintainability, supportability, and operability are no longer effects of design; they have moved to the forefront and are affecting design. A primary focus of this shift has been a planned decrease in vehicle turnaround time. Potentials for instituting this decrease include attacking the issues of removing, refurbishing, and replacing the engines after each flight. less, it is important to understand the operational affects of an engine on turnaround time, ground support personnel and equipment. One tool for visualizing this relationship involves the creation of a Discrete Event Simulation (DES). A DES model can be used to run a series of trade studies to determine if the engine is meeting its requirements, and, if not, what can be altered to bring it into compliance. Using DES, it is possible to look at the ways in which labor requirements, parallel maintenance versus serial maintenance, and maintenance scheduling affect the overall turnaround time. A detailed DES model of the Space Shuttle Main Engines (SSME) has been developed. Trades may be performed using the SSME Processing Model to see where maintenance bottlenecks occur, what the benefits (if any) are of increasing the numbers of personnel, or the number and location of facilities, in addition to trades previously mentioned, all with the goal of optimizing the operational turnaround time and minimizing operational cost. The SSME Processing Model was developed in such a way that it can easily be used as a foundation for developing DES models of other operational or developmental reusable engines. Performing a DES on a developmental engine during the conceptual phase makes it easier to affect the design and make changes to bring about a decrease in turnaround time and costs.

Nix, Michael↗