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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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The use of new technology and advanced emission control hardware to reduce emissions from the direct combustion of coal to produce electricity in California is considered. The technical feasibilty of a demonstration project on an existing 81-MW boiler is demonstrated.
The basic Space Shuttle Main Engine (SSME) concept called for an operational life of 55 missions, as well as an exceptionally high specific impulse. While these requirements are considered achievable in the cases of the basic engine structure, powerhead, main combustion chamber, propellant ducts, hydraulic lines and engine controller, other components may experience excessive wear and deterioration and may not achieve specification life without removal and overhaul. Most SSME components have been designed as line replaceable units, which have demonstrated good operational durability and will meet flight certification requirement upgrading. The Flight Confidence Program of the SSME test program currently being instituted will recertify existing engines for use in more than the 10 flights stipulated until now. The testing of two engines in repeated cycles should add 10 flights for every two successfully completed certification cycles.
The altitude operational characteristics of the prototype J40-WE-8 turbojet engine were determined in the NACA Lewis altitude wind tunnel. The operational characteristics that were obtained include (1) compressor surge characteristics, (2) acceleration and deceleration rates, (3) steady-state windmilling-engine characteristics, (4) altitude ignition characteristics, and (5) the effect of two grades of fuel on steady-state engine performance at an altitude of 45,000 feet and a flight Mach number of 0.20. The altitude ignition characteristics determined include the effects of (1) ignition fuel flow, (2) two ignition procedures, (3) fuel temperatures, and (4) two grades of fuel on ignition characteristics. The compressor surge line, when presented as a function of compressor pressure ratio and corrected engine speed, was not affected by changes in flight condition and was independent of engine-inlet installation and of the manner in which surge was approached, rapidly or slowly. Also there was no effect of altitude or engine-inlet installation on the compressor surge recovery line when presented as a function of compressor pressure ratio and corrected engine speed.
The Advanced Turbine Technologies Application Project (ATTAP) is in the fifth year of a multiyear development program to bring the automotive gas turbine engine to a state at which industry can make commercialization decisions. Activities during the past year included reference powertrain design updates, test-bed engine design and development, ceramic component design, materials and component characterization, ceramic component process development and fabrication, ceramic component rig testing, and test-bed engine fabrication and testing. Engine design and development included mechanical design, combustion system development, alternate aerodynamic flow testing, and controls development. Design activities included development of the ceramic gasifier turbine static structure, the ceramic gasifier rotor, and the ceramic power turbine rotor. Material characterization efforts included the testing and evaluation of five candidate high temperature ceramic materials. Ceramic component process development and fabrication, with the objective of approaching automotive volumes and costs, continued for the gasifier turbine rotor, gasifier turbine scroll, extruded regenerator disks, and thermal insulation. Engine and rig fabrication, testing, and development supported improvements in ceramic component technology. Total test time in 1992 amounted to 599 hours, of which 147 hours were engine testing and 452 were hot rig testing.
A research program is currently underway with the final objective of developing a fundamental understanding of the controlling mechanisms underlying the flammability diagrams of solid combustible materials and their derived fire properties. Given that there is a high possibility of an accidental fire occurring in a space-based facility, understanding the fire properties of materials that will be used in such facilities is of critical importance. With this purpose, the flammability diagrams of the materials, as those produced by the Lateral Ignition and Flame Spread Test (LIFT) apparatus and by a new forced flow device, the Forced Flow Ignition and Flame Spread Test (FIST) apparatus, will be obtained. The specific objective of the program is to apply the new flammability apparatus, which will more accurately reflect the potential ambient conditions of space-based environments, to the characterization of the materials for space applications. This paper presents a parametric study of oxidizer flow effects on the ignition curve of the flammability diagrams of PMMA. The dependence of the ignition delay time on the external radiant flux and either the sample width (LIFT) or the flow velocity (FIST) has been studied. Although preliminary, the results indicate that natural and forced convection flow changes, affect the characteristics of the ignition curves of the flammability diagrams. The major effect on the ignition time appears to be due to convective transfer variations at the fuel surface. At high radiant fluxes or high flow velocities, however, it appears that gas phase processes become increasingly important, affecting the overall ignition delay time. A numerical analysis of the solid fuel heating and pyrolysis has also been developed. The theoretical predictions approximate the experiments well for conditions in which the gas phase induction time is negligible.
Liquid oxygen/liquid methane (LO2/LCH4 ) has recently been viewed as a potential green propulsion system for both the Altair ascent main engine (AME) and reaction control system (RCS). The Propulsion and Cryogenic Advanced Development Project (PCAD) has been tasked by NASA to develop these green propellant systems to enable safe and cost effective exploration missions. However, experience with LO2/LCH4 as a propellant combination is limited, so testing of these systems is critical to demonstrating reliable ignition and performance. A test program of a 100 lb f reaction control engine (RCE) is underway at the Altitude Combustion Stand (ACS) of the NASA Glenn Research Center, with a focus on conducting tests at altitude conditions. These tests include a unique propellant conditioning feed system (PCFS) which allows for the inlet conditions of the propellant to be varied to test warm to subcooled liquid propellant temperatures. Engine performance, including thrust, c* and vacuum specific impulse (I(sub sp,vac)) will be presented as a function of propellant temperature conditions. In general, the engine performed as expected, with higher performance at warmer propellant temperatures but better efficiency at lower propellant temperatures. Mixture ratio effects were inconclusive within the uncertainty bands of data, but qualitatively showed higher performance at lower ratios.
Topics include: Relative-Motion Sensors and Actuators for Two Optical Tables; Improved Position Sensor for Feedback Control of Levitation; Compact Tactile Sensors for Robot Fingers; Improved Ion-Channel Biosensors; Suspended-Patch Antenna With Inverted, EM-Coupled Feed; System Would Predictively Preempt Traffic Lights for Emergency Vehicles; Optical Position Encoders for High or Low Temperatures; Inter-Valence-Subband/Conduction-Band-Transport IR Detectors; Additional Drive Circuitry for Piezoelectric Screw Motors; Software for Use with Optoelectronic Measuring Tool; Coordinating Shared Activities; Software Reduces Radio-Interference Effects in Radar Data; Using Iron to Treat Chlorohydrocarbon-Contaminated Soil; Thermally Insulating, Kinematic Tensioned-Fiber Suspension; Back Actuators for Segmented Mirrors and Other Applications; Mechanism for Self-Reacted Friction Stir Welding; Lightweight Exoskeletons with Controllable Actuators; Miniature Robotic Submarine for Exploring Harsh Environments; Electron-Spin Filters Based on the Rashba Effect; Diffusion-Cooled Tantalum Hot-Electron Bolometer Mixers; Tunable Optical True-Time Delay Devices Would Exploit EIT; Fast Query-Optimized Kernel-Machine Classification; Indentured Parts List Maintenance and Part Assembly Capture Tool - IMPACT; An Architecture for Controlling Multiple Robots; Progress in Fabrication of Rocket Combustion Chambers by VPS; CHEM-Based Self-Deploying Spacecraft Radar Antennas; Scalable Multiprocessor for High-Speed Computing in Space; and Simple Systems for Detecting Spacecraft Meteoroid Punctures.
This paper presents a numerical model of pressurization of a cryogenic propellant tank for the Integrated Vehicle Fluid (IVF) system using the Generalized Fluid System Simulation Program (GFSSP). The IVF propulsion system, being developed by United Launch Alliance, uses boiloff propellants to drive thrusters for the reaction control system as well as to run internal combustion engines to develop power and drive compressors to pressurize propellant tanks. NASA Marshall Space Flight Center (MSFC) has been running tests to verify the functioning of the IVF system using a flight tank. GFSSP, a finite volume based flow network analysis software developed at MSFC, has been used to develop an integrated model of the tank and the pressurization system. This paper presents an iterative algorithm for converging the interface boundary conditions between different component models of a large system model. The model results have been compared with test data.
Retreating blade stall is a well-known phenomenon that limits rotorcraft speed, maneuverability, and efficiency. Airfoil dynamic stall is a simpler problem, which demonstrates many of the same flow phenomena. Combustion Powered Actuation (COMPACT) is an active flow control technology, which at the outset of this work, had been shown to mitigate static and dynamic stall at low Mach numbers. The attributes of this technology suggested strong potential for success at higher Mach numbers, but such experiments had never been conducted. The work detailed in this report documents a 3-year effort focused on assessing the effectiveness of COMPACT for dynamic stall suppression at freestream conditions up to Mach 0.5. The work done has focused on implementing COMPACT on a high-lift rotorcraft airfoil: the VR-12. This selection was made in order to ensure that any measured benefits are over and above the capabilities of state-of-the-art high-lift rotorcraft airfoils. The detailed Computational Fluid Dynamics (CFD) simulations, wind-tunnel experiments, and system-level modeling conducted have shown the following: (1) COMPACT, in its current state of development, is capable of reducing the adverse effects of deep dynamic stall at Mach numbers up to 0.4; (2) The two-dimensional (2D) CFD results trend well compared to the experiments; and (3) Implementation of the CFD results into a system-level model suggest that significant rotor-level benefits are possible.
This paper presents a numerical model of the pressurization loop of the Integrated Vehicle Fluids (IVF) system using the Generalized Fluid System Simulation Program (GFSSP). The IVF propulsion system, being developed by United Launch Alliance to reduce system weight and enhance reliability, uses boiloff propellants to drive thrusters for the reaction control system as well as to run internal combustion engines to develop power and drive compressors to pressurize propellant tanks. NASA Marshall Space Flight Center (MSFC) conducted tests to verify the functioning of the IVF system using a flight-like tank. GFSSP, a finite volume based flow network analysis software developed at MSFC, has been used to support the test program. This paper presents the simulation of three different test series, comparison of numerical prediction and test data and a novel method of presenting data in a dimensionless form. The paper also presents a methodology of implementing a compressor map in a system level code.
The Lewis Research Center is developing a series of microgravity combustion experiments for the Space Shuttle. The Solid Surface Combustion Experiment (SSCE) is the first to be completed. SSCE will study flame spreading over thermally thin fuels (ashless filter paper) under microgravity conditions. The flight hardware consists of a combustion chamber containing the sample and a computer which takes the data and controls the experiment. Experimental data will include gas-phase and solid-phase temperature measurements and motion pictures of the combustion process. Flame spread rates will be determined from the motion pictures.
The Lewis Research Center is developing a series of microgravity combustion experiments for the Space Shuttle. The Solid Surface Combustion Experiment (SSCE) is the first to be completed. SSCE will study flame spreading over thermally thin fuels (ashless filter paper) under microgravity conditions. The flight hardware consists of a combustion chamber containing the sample and a computer which takes the data and controls the experiment. Experimental data will include gas-phase and solid-phase temperature measurements and motion pictures of the combustion process. Flame spread rates will be determined from the motion pictures.
Pulsed combustion appears to have the potential to provide for rapid fuel/air mixing, compact and economical combustors, and reduced exhaust emissions. The objective of this Flight-Definition experiment (PuFF, for Pulsed-Fully Flames) is to increase the fundamental understanding of the fuel/air mixing and combustion behavior of pulsed, turbulent diffusion flames by conducting experiments in microgravity. In this research the fuel jet is fully modulated (i.e., completely shut off between pulses) by an externally controlled valve system. This gives rise to drastic modification of the combustion and flow characteristics of flames, leading to enhanced fuel/air mixing mechanisms not operative for the case of acoustically excited or partially-modulated jets. The fully-modulated injection approach also simplifies the combustion process by avoiding the acoustic forcing generally present in pulsed combustors. Relatively little is known about the behavior of turbulent flames in reduced-gravity conditions, even in the absence of pulsing. Fundamental issues addressed in this experiment include the impact of buoyancy on the fuel/air mixing and combustion characteristics of fully-modulated flames. It is also important for the planned space experiments to establish the effects of confinement and oxidizer co-flow on these flames.
Reduced gravity is used to carry out scientific investigations of droplet combustion. In earlier work a preliminary conceptual design had been developed for droplet burning experiments for Spacelab. Refinements of the earlier work with special consideration given to possible experiments for Mid-Deck Modules of the Space Shuttle is discussed. A re-evaluation of suitable experiments on droplet combustion is done to ascertain whether influences of reduced buoyancy on time-dependent processes of heat and mass transfer in the gas or in the liquid on extinction processes or on disruptive burning phenomena are best suited for further investigation. Components of the experimental apparatus, which include a droplet dispensing system, a droplet positioning system, a droplet ignition system and provision for recording, primarily photographically, the combustion of the individual droplet in a chamber having a controlled gas atmosphere, are studied to determine optimal approaches to the experimental design. Methods for data reduction and interpretation are being made more specific, particularly in the context of an objective to calculate overall gas-phase chemical-kinetic parameters for the combustion from observations of extinction conditions in different atmospheres.
Pulsed combustion appears to have the potential to provide for rapid fuel/air mixing, compact and economical combustors, and reduced exhaust emissions. The objective of this experiment (PuFF, for Pulsed-Fully Flames) is to increase the fundamental understanding of the fuel/air mixing and combustion behavior of pulsed, turbulent diffusion flames by conducting experiments in microgravity. In this research the fuel jet is fully-modulated (i.e., completely shut off between pulses) by an externally controlled valve system. This gives rise to drastic modification of the combustion and flow characteristics of flames, leading to enhanced fuel/air mixing compared to acoustically excited or partially-modulated jets. Normal-gravity experiments suggest that the fully-modulated technique also has the potential for producing turbulent jet flames significantly more compact than steady flames with no increase in exhaust emissions. The technique also simplifies the combustion process by avoiding the acoustic forcing generally present in pulsed combustors. Fundamental issues addressed in this experiment include the impact of buoyancy on the structure and flame length, temperatures, radiation, and emissions of fully-modulated flames.
As a result of the ongoing exploration of Mars and the several unmanned and possibly manned missions planned for the near future, increased attention has been given to the use of the natural resources of the planet for rocket propellant production and energy generation. Since the atmosphere of Mars consists of approximately 95% carbon dioxide (CO2), this gas is the resource of choice to be employed for these purposes. Since many metals burn vigorously with CO2, these may be used as an energy source or as propellants for a research vehicle on the surface of Mars. Shafirovich and Goldshleger conducted experiments with spherical particles up to 2.5 mm in diameter and found that the burning process was controlled by diffusion and that the particles exhibited pulsating combustion due to superheating of the Mg vapor trapped inside a protective oxide shell. They also proposed a reaction mechanism based on the gas-phase reaction, Mg + CO2 yields MgO + CO and the heterogeneous reaction Mg + CO yields MgO + C occurring on the sample surface. In all the above studies with large Mg particles, the burning process is invariably influenced by strong convective currents that accelerate the combustion reaction and shorten the burning times. Although these currents are nearly absent in the burning of small particles, the high emissivity of the flames, rapid reaction, and small length scales make the gathering of any useful information on burning rates and flame structure very difficult. The goal of this investigation is to provide a detailed study of flame structure by taking advantage of large, free-floating spherical metal samples and their corresponding long burning times available in a weightless environment. The use of reduced gravity is essential to eliminate the intrusive buoyant flows that plague high temperature metal reactions, to remove the destructive effect of gravity on the shape of molten metal samples, and to study the combustion behavior of metals in the presence of solid oxides undisturbed by natural convection. This work presents the most complete modeling of metal particle burning to date for Mg with CO2 and O2.
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.
Active Control can help to meet future engine requirements by an active improvement of the component characteristics. The concept is based on an intelligent control logic, which senses actual operating conditions and reacts with adequate actuator action. This approach can directly improve engine characteristics as performance, operability, durability and emissions on the one hand. On the other hand active control addresses the design constrains imposed by unsteady phenomena like inlet distortion, compressor surge, combustion instability, flow separations, vibration and noise, which only occur during exceptional operating conditions. The feasibility and effectiveness of active control technologies have been demonstrated in lab-scale tests. This chapter describes a broad range of promising applications for each engine component. Significant efforts in research and development remain to implement these technologies in engine rig and finally production engines and to demonstrate today s engine generation airworthiness, safety, reliability, and durability requirements. Active control applications are in particular limited by the gap between available and advanced sensors and actuators, which allow an operation in the harsh environment in an aero engine. The operating and performance requirements for actuators and sensors are outlined for each of the gas turbine sections from inlet to nozzle.