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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 181 records · Page 10

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NASA NTRS1989

Orbit transfer rocket engine technology program: Oxygen materials compatibility testing

Particle impact and frictional heating tests of metals in high pressure oxygen, are conducted in support of the design of an advanced rocket engine oxygen turbopump. Materials having a wide range of thermodynamic properties including heat of combustion and thermal diffusivity were compared in their resistance to ignition and sustained burning. Copper, nickel and their alloys were found superior to iron based and stainless steel alloys. Some materials became more difficult to ignite as oxygen pressure was increased from 7 to 21 MPa (1000 to 3000 psia).

Schoenman, Leonard↗

NASA NTRS1997

Oxygen Compatibility Screening Tests in Oxygen-Rich Combustion Environment

The identification and characterization of oxygen-rich compatible materials enables full-flow, staged combustion designs. Although these oxygen-rich designs offer significant cost, performance, and reliability benefits over existing systems, they have never been used operationally by the United States. If these systems are to be realized, it is critical to understand the long-term oxidative stability in high-temperature, high-pressure, oxygen-rich combustion environments. A unique facility has been constructed at the NASA Lewis Research Center to conduct tests of small-scale rocket engine materials and subcomponents in an oxygen-rich combustion environment that closely approximates a full-scale rocket engine. Thus, a broad range of advanced materials and concepts can be screened in a timely manner and at a relatively low cost.

Eckel, Anerew J.↗

NASA NTRS2012

Experimental Investigation of Augmented Spark Ignition of a LO2/LCH4 Reaction Control Engine at Altitude Conditions

The use of nontoxic propellants in future exploration vehicles would enable safer, more cost-effective mission scenarios. One promising green alternative to existing hypergols is liquid methane (LCH4) with liquid oxygen (LO2). A 100 lbf LO2/LCH4 engine was developed under the NASA Propulsion and Cryogenic Advanced Development project and tested at the NASA Glenn Research Center Altitude Combustion Stand in a low pressure environment. High ignition energy is a perceived drawback of this propellant combination; so this ignition margin test program examined ignition performance versus delivered spark energy. Sensitivity of ignition to spark timing and repetition rate was also explored. Three different exciter units were used with the engine s augmented (torch) igniter. Captured waveforms indicated spark behavior in hot fire conditions was inconsistent compared to the well-behaved dry sparks. This suggests that rising pressure and flow rate increase spark impedance and may at some point compromise an exciter s ability to complete each spark. The reduced spark energies of such quenched deliveries resulted in more erratic ignitions, decreasing ignition probability. The timing of the sparks relative to the pressure/flow conditions also impacted the probability of ignition. Sparks occurring early in the flow could trigger ignition with energies as low as 1 to 6 mJ, though multiple, similarly timed sparks of 55 to 75 mJ were required for reliable ignition. Delayed spark application and reduced spark repetition rate both correlated with late and occasional failed ignitions. An optimum time interval for spark application and ignition therefore coincides with propellant introduction to the igniter.

Kleinhenz, Julie↗

NASA NTRS1993

Design and test of a small two stage counter-rotating turbine for rocket engine application

The aerodynamic design and rig test evaluation of a small counter-rotating turbine system is described. The technology represented by this turbine is being developed for application in an advanced upper stage rocket engine turbopump. This engine will employ an oxygen/hydrogen expander cycle and achieve high performance through efficient combustion, high combustion pressure, and high area ratio exhaust nozzle expansion. Engine performance goals require that the turbopump drive turbines achieve high efficiency at low gas flow rates. The low flow rates result in very small airfoil diameter, height and chord. The high efficiency and small size requirements present a challenging turbine design problem. The unconventional approach employed to meet this challenge is described, along with the detailed design process and resulting airfoil configurations. The method and results of full scale aerodynamic performance evaluation testing of both one and two stage configurations, as well as operation without the secondary stage stator are presented. The overall results of this effort illustrate that advanced aerodynamic design tools and hardware fabrication techniques have provided improved capability to produce small high performance turbines for advanced rocket engines.

Huber, F. W.↗

NASA NTRS1999

NASA GRC's High Pressure Burner Rig Facility and Materials Test Capabilities

The High Pressure Burner Rig (HPBR) at NASA Glenn Research Center is a high-velocity. pressurized combustion test rig used for high-temperature environmental durability studies of advanced materials and components. The facility burns jet fuel and air in controlled ratios, simulating combustion gas chemistries and temperatures that are realistic to those in gas turbine engines. In addition, the test section is capable of simulating the pressures and gas velocities representative of today's aircraft. The HPBR provides a relatively inexpensive. yet sophisticated means for researchers to study the high-temperature oxidation of advanced materials. The facility has the unique capability of operating under both fuel-lean and fuel-rich gas mixtures. using a fume incinerator to eliminate any harmful byproduct emissions (CO, H2S) of rich-burn operation. Test samples are easily accessible for ongoing inspection and documentation of weight change, thickness, cracking, and other metrics. Temperature measurement is available in the form of both thermocouples and optical pyrometery. and the facility is equipped with quartz windows for observation and video taping. Operating conditions include: (1) 1.0 kg/sec (2.0 lbm/sec) combustion and secondary cooling airflow capability: (2) Equivalence ratios of 0.5- 1.0 (lean) to 1.5-2.0 (rich), with typically 10% H2O vapor pressure: (3) Gas temperatures ranging 700-1650 C (1300-3000 F): (4) Test pressures ranging 4-12 atmospheres: (5) Gas flow velocities ranging 10-30 m/s (50-100) ft/sec.: and (6) Cyclic and steady-state exposure capabilities. The facility has historically been used to test coupon-size materials. including metals and ceramics. However complex-shaped components have also been tested including cylinders, airfoils, and film-cooled end walls. The facility has also been used to develop thin-film temperature measurement sensors.

Robinson, R. Craig↗

NASA NTRS1979

Durability testing of advanced catalysts and catalyst supports for gas turbine engine combustors

The paper presents new information on the durability of a CATCOM catalyst operating at low-emission combustion temperatures (about 1527 K) with a liquid fuel, No. 2 diesel. Information on the activity of No. 2 diesel after 1000 hr of aging is given. In addition, a unique in situ activity test developed for monitoring the subtle changes in the catalyst activity of the CATCOM catalyst is also detailed. The study demonstrated the feasibility of using a CATCOM catalyst in catalytically supported thermal combustion for extended operating periods

Heck, R. M.↗

NASA NTRS1980

Advanced Technology Spark-Ignition Aircraft Piston Engine Design Study

The advanced technology, spark ignition, aircraft piston engine design study was conducted to determine the improvements that could be made by taking advantage of technology that could reasonably be expected to be made available for an engine intended for production by January 1, 1990. Two engines were proposed to account for levels of technology considered to be moderate risk and high risk. The moderate risk technology engine is a homogeneous charge engine operating on avgas and offers a 40% improvement in transportation efficiency over present designs. The high risk technology engine, with a stratified charge combustion system using kerosene-based jet fuel, projects a 65% improvement in transportation efficiency. Technology enablement program plans are proposed herein to set a timetable for the successful integration of each item of required advanced technology into the engine design.

Stuckas, K. J.↗

NASA NTRS1983

The spark-ignition aircraft piston engine of the future

The advanced technology, spark ignition, aircraft piston engine design study was conducted to determine the improvements that could be made by taking advantage of technology that could reasonably be expected to be made available for an engine intended for production by January 1, 1990. Two engines were proposed to account for levels of technology considered to be moderate risk and high risk. The moderate risk technology engine is a homogeneous charge engine operating on avgas and offers a 40% improvement in transportation efficiency over present designs. The high risk technology engine, with a stratified charge combustion system using kerosene-based jet fuel, projects a 65% improvement in transportation efficiency. Technology enablement program plans are proposed herein to set a timetable for the successful integration of each time of required advanced technology into the engine design.

Stuckas, K. J.↗

NASA NTRS1990

Rocket combustion chamber life-enhancing design concepts

NASA continues to pursue technologies which can lead to an increase in life and reduce the costs of fabrication of the Space Shuttle Main Engine. The joint NASA/Air Force Advanced Launch System Program has set its prime objectives to be high reliability and low cost for their new advanced booster engine. In order to meet these objectives, NASA will utilize the results of several ongoing programs to provide the required technologies. An overview is presented of those programs which address life enhancing design concepts for the combustion chamber. Seven different design concepts, which reduce the thermal strain and/or increase the material strength of the combustion chamber liner wall are discussed. Subscale rocket test results are presented, where available, for life enhancing design concepts. Two techniques for reducing chamber fabrication costs are discussed, as well as issues relating to hydrocarbon fuels/combustion chamber liner materials compatibility.

Quentmeyer, Richard J.↗

NASA NTRS1990

Rocket combustion chamber life-enhancing design concepts

Quentmeyer, R. J.↗

NASA NTRS1987

A variable geometry combustor for broadened properties fuels

A program was conducted to design and develop a variable geometry combustor, sized for the cycle and envelope of a large commercial turbofan engine. The combustor uses a variable area swirl cup to control stoichiometry in the primary combustion zone. Potential advantages of this design include improved capability to burn non-standard fuels, short system length, and increased operating temperature range for advanced high performance engine cycles. After considerable development, key program emissons and performance goals were met with the variable geometry combustor. Primary development efforts were to evolve improved variable swirl cup configurations. In particular, air leakage through the variable area swirl cup had a strong effect on low power emissions and performance, while smoke level at high power was affected by features for improved mixing of the fuel and swirler air flow. Additional design and development is still needed to evolve a practical variable geometry combustor.

Dodds, W. J.↗

NASA NTRS1984

Research on aviation fuel instability

Current aircraft turbine fuels do not present a significant problem with fuel thermal stability. However, turbine fuels with broadened properties or nonpetroleum derived fuels may have reduced thermal stability because of their higher content of olefins, heteroatoms, and trace metals. Moreover, advanced turbine engines will increase the thermal stress on fuels because of their higher pressure ratios and combustion temperature. In recognition of the importance of this problem, NASA Lewis is currently engaged in a broadly based research effort to better understand the underlying causes of fuel thermal degradation. The progress and status of our various activities in this area are discussed. Topics covered include: nature of fuel instability and its temperature dependence, methods of measuring the instability, chemical mechanisms involved in deposit formation, and instrumental methods for characterizing fuel deposits. Finally, some preliminary thoughts on design approaches for minimizing the effects of lowered thermal stability are briefly discussed.

Baker, C. E.↗

NASA NTRS1992

A new unsteady mixing model to predict NO(x) production during rapid mixing in a dual-stage combustor

An advanced gas turbine engine to power supersonic transport aircraft is currently under study. In addition to high combustion efficiency requirements, environmental concerns have placed stringent restrictions on the pollutant emissions from these engines. A combustor design with the potential for minimizing pollutants such as NO(x) emissions is undergoing experimental evaluation. A major technical issue in the design of this combustor is how to rapidly mix the hot, fuel-rich primary zone product with the secondary diluent air to obtain a fuel-lean mixture for combustion in the second stage. Numerical predictions using steady-state methods cannot account for the unsteady phenomena in the mixing region. Therefore, to evaluate the effect of unsteady mixing and combustion processes, a novel unsteady mixing model is demonstrated here. This model has been used to study multispecies mixing as well as propane-air and hydrogen-air jet nonpremixed flames, and has been used to predict NO(x) production in the mixing region. Comparison with available experimental data show good agreement, thereby providing validation of the mixing model. With this demonstration, this mixing model is ready to be implemented in conjunction with steady-state prediction methods and provide an improved engineering design analysis tool.

Menon, Suresh↗

NASA NTRS1992

A new unsteady mixing model to predict NO(x) production during rapid mixing in a dual-stage combustor

An advanced gas turbine engine to power supersonic transport aircraft is currently under study. In addition to high combustion efficiency requirements, environmental concerns have placed stringent restrictions on the pollutant emissions from these engines. A dual-stage combustor with the potential for minimizing pollutants such as NO(x) emissions is undergoing experimental evaluation. A major technical issue in the design of this combustor is how to rapidly mix the hot, fuel-rich primary stage product with the secondary diluent air to obtain a fuel-lean mixture for combustion in the secondary stage. Numerical design studies using steady-state methods cannot account for the unsteady phenomena in the mixing region. Therefore, to evaluate the effect of unsteady mixing and combustion processes, a novel unsteady mixing model is demonstrated here. This model has been used in a stand-alone mode to study mixing and combustion in hydrogen-air nonpremixed jet flames. NO(x) production in these jet flames was also predicted. Comparison of the computed results with experimental data show good agreement thereby providing validation of the mixing model.

Menon, Suresh↗

NASA NTRS2005

Model-Based, Multiscale Self-Tuning Controller Developed for Active Combustion Control

New challenges concerning system health-monitoring and life-extending robust controls for the Ultra-Efficient Engine Technology Project, as well as other advanced engine and power system concepts at NASA and elsewhere, have renewed the control community s interest in smart, model-based methods. In particular, these challenges have further motivated efforts at the NASA Glenn Research Center to exploit the versatility and superiority of the dynamic features extraction of multiscale analysis for controls--such as with "wavelets" and "wavelet filter-banks.' The accomplishments reported herein pertain to the active suppression of combustion instabilities in liquid-fuel combustors via fuel modulation. The fundamentals and initial success of this innovation were reported for a unique demonstration of active combustion control (a research collaboration of NASA Glenn with Pratt & Whitney and the United Technologies Research Center, UTRC). This demonstration, conducted in 2002 at UTRC on the NASA single nozzle rig (SNR) combustor, was the first known suppression of high-frequency instability with a liquid-fueled combustor. The SNR is based on a high-powered military engine combustor that exhibited well-known instabilities.

Le, Dzu K.↗

NASA NTRS2006

NASA's J-2X Engine Builds on the Apollo Program for Lunar Return Missions

In January 2006, NASA streamlined its U.S. Vision for Space Exploration hardware development approach for replacing the Space Shuttle after it is retired in 2010. The revised CLV upper stage will use the J-2X engine, a derivative of NASA s Apollo Program Saturn V s S-II and S-IVB main propulsion, which will also serve as the Earth Departure Stage (EDS) engine. This paper gives details of how the J- 2X engine effort mitigates risk by building on the Apollo Program and other lessons learned to deliver a human-rated engine that is on an aggressive development schedule, with first demonstration flight in 2010 and human test flights in 2012. It is well documented that propulsion is historically a high-risk area. NASA s risk reduction strategy for the J-2X engine design, development, test, and evaluation is to build upon heritage hardware and apply valuable experience gained from past development efforts. In addition, NASA and its industry partner, Rocketdyne, which originally built the J-2, have tapped into their extensive databases and are applying lessons conveyed firsthand by Apollo-era veterans of America s first round of Moon missions in the 1960s and 1970s. NASA s development approach for the J-2X engine includes early requirements definition and management; designing-in lessons learned from the 5-2 heritage programs; initiating long-lead procurement items before Preliminary Desi& Review; incorporating design features for anticipated EDS requirements; identifying facilities for sea-level and altitude testing; and starting ground support equipment and logistics planning at an early stage. Other risk reduction strategies include utilizing a proven gas generator cycle with recent development experience; utilizing existing turbomachinery ; applying current and recent main combustion chamber (Integrated Powerhead Demonstrator) and channel wall nozzle (COBRA) advances; and performing rigorous development, qualification, and certification testing of the engine system, with a philosophy of "test what you fly, and fly what you test". These and other active risk management strategies are in place to deliver the J-2X engine for LEO and lunar return missions as outlined in the U.S. Vision for Space Exploration.

Snoddy, Jimmy R.↗

NASA NTRS1999

NCC: A Multidisciplinary Design/Analysis Tool for Combustion Systems

Liu, Nan-Suey↗

NASA NTRS2001

Overview of the NCC

A multi-disciplinary design/analysis tool for combustion systems is critical for optimizing the low-emission, high-performance combustor design process. Based on discussions between then NASA Lewis Research Center and the jet engine companies, an industry-government team was formed in early 1995 to develop the National Combustion Code (NCC), which is an integrated system of computer codes for the design and analysis of combustion systems. NCC has advanced features that address the need to meet designer's requirements such as "assured accuracy", "fast turnaround", and "acceptable cost". The NCC development team is comprised of Allison Engine Company (Allison), CFD Research Corporation (CFDRC), GE Aircraft Engines (GEAE), NASA Glenn Research Center (LeRC), and Pratt & Whitney (P&W). The "unstructured mesh" capability and "parallel computing" are fundamental features of NCC from its inception. The NCC system is composed of a set of "elements" which includes grid generator, main flow solver, turbulence module, turbulence and chemistry interaction module, chemistry module, spray module, radiation heat transfer module, data visualization module, and a post-processor for evaluating engine performance parameters. Each element may have contributions from several team members. Such a multi-source multi-element system needs to be integrated in a way that facilitates inter-module data communication, flexibility in module selection, and ease of integration. The development of the NCC beta version was essentially completed in June 1998. Technical details of the NCC elements are given in the Reference List. Elements such as the baseline flow solver, turbulence module, and the chemistry module, have been extensively validated; and their parallel performance on large-scale parallel systems has been evaluated and optimized. However the scalar PDF module and the Spray module, as well as their coupling with the baseline flow solver, were developed in a small-scale distributed computing environment. As a result, the validation of the NCC beta version as a whole was quite limited. Current effort has been focused on the validation of the integrated code and the evaluation/optimization of its overall performance on large-scale parallel systems.

Liu, Nan-Suey↗