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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 361 records · Page 20
All sourcesNASA NTRSDOE OSTI
NASA NTRS1922

The Effect of Longitudinal Moment of Inertia upon Dynamic Stability

Free flight tests were carried out to show whether the longitudinal oscillations of a standard S.E.5A airplane are noticeably affected if its longitudinal moment of inertia is increased. These oscillations were taken by means of a self-recording instrument, the airplane having first its ordinary moment of inertia and then one increased by 14 percent. The period of oscillation was slightly longer after the increase of the moment of inertia, but the damping was not affected. Presented here are test results from an investigation to determine the relative performance of a single-cylinder, high-speed, compression-ignition engine when using fuel injection valve nozzles with different numbers, sizes, and directions of round orifices. A spring loaded, automatic injection valve was used. It was centrally located at the top of a vertical disk-type combustion chamber formed between horizontally opposed inlet and exhaust valves of a 5-inch by 7-inch engine. A series of fuel injection valve nozzles with different arrangements of round orifices were tested, starting with orifices so small that impingement on the combustion chamber walls was impossible and increasing beyond the start of impingement. a :A table and curves are presented showing the performance of the engine with different nozzles. The test results are discussed, and some probable reasons given for the variation in performance with different nozzles on the basis of spray distribution.

MOMENTS OF INERTIA - ROYAL AIRCRAFT FACTORY S E 5A↗
NASA NTRS2003

System Identification of X-33 Neural Network

Modern flight control research has improved spacecraft survivability as its goal. To this end we need to have a failure detection system on board. In case the spacecraft is performing imperfectly, reconfiguration of control is needed. For that purpose we need to have parameter identification of spacecraft dynamics. Parameter identification of a system is called system identification. We treat the system as a black box which receives some inputs that lead to some outputs. The question is: what kind of parameters for a particular black box can correlate the observed inputs and outputs? Can these parameters help us to predict the outputs for a new given set of inputs? This is the basic problem of system identification. The X33 was supposed to have the onboard capability of evaluating the current performance and if needed to take the corrective measures to adapt to desired performance. The X33 is comprised of both rocket and aircraft vehicle design characteristics and requires, in general, analytical methods for evaluating its flight performance. Its flight consists of four phases: ascent, transition, entry and TAEM (Terminal Area Energy Management). It spends about 200 seconds in ascent phase, reaching an altitude of about 180,000 feet and a speed of about 10 to 15 Mach. During the transition phase which lasts only about 30 seconds, its altitude may increase to about 190,000 feet but its speed is reduced to about 9 Mach. At the beginning of this phase, the Main Engine is Cut Off (MECO) and the control is reconfigured with the help of aerosurfaces (four elevons, two flaps and two rudders) and reaction control system (RCS). The entry phase brings down the altitude of X33 to about 90,000 feet and its speed to about Mach 3. It spends about 250 seconds in this phase. Main engine is still cut off and the vehicle is controlled by complex maneuvers of aerosurfaces. The last phase TAEM lasts for about 450 seconds and the altitude and speed, both are reduced to zero. The present attempt, as a start, focuses only on the entry phase. Since the main engine remains cut off in this phase, there is no thrust acting on the system. This considerably simplifies the equations of motion. We introduce another simplification by assuming the system to be linear after some non-linearities are removed analytically from our consideration. Under these assumptions, the problem could be solved by Classical Statistics by employing the least sum of squares approach. Instead we chose to use the Neural Network method. This method has many advantages. It is modern, more efficient, can be adapted to work even when the assumptions are diluted. In fact, Neural Networks try to model the human brain and are capable of pattern recognition.

Aggarwal, Shiv↗
NASA NTRS1975

Quiet clean short-haul experimental engine /QCSEE/ design rationale

The principal design features of the NASA QCSEE Under-The-Wing and Over-The-Wing powered lift propulsion systems are given. In the UTW engine, these include noise reduction features, a variable pitch low pressure ratio fan, a fan drive reduction gear, an advanced core and low pressure turbine with a low pollution combustor, a digital control, and advanced composite construction for the inlet, fan frame, fan exhaust duct, and variable area fan exhaust nozzle. The OTW engine is similar but has higher fan pressure and a fixed pitch fan. Both engines are scheduled to be fabricated and tested starting in 1976.

Adamson, A. P.↗
NASA NTRS1976

Preliminary compressor design study for an advanced multistage axial flow compressor

An optimum, axial flow, high pressure ratio compressor for a turbofan engine was defined for commercial subsonic transport service starting in the late 1980's. Projected 1985 technologies were used and applied to compressors with an 18:1 pressure ratio having 6 to 12 stages. A matrix of 49 compressors was developed by statistical techniques. The compressors were evaluated by means of computer programs in terms of various airline economic figures of merit such as return on investment and direct-operating cost. The optimum configuration was determined to be a high speed, 8-stage compressor with an average blading aspect ratio of 1.15.

Marman, H. V.↗
NASA NTRS1977

Experimental clean combustor program, phase 3

A two-stage vortex burning and mixing combustor and associated fuel system components were successfully tested at steady state and transient operating conditions. The combustor exceeded the program goals for all three emissions species, with oxides of nitrogen 10 percent below the goal, carbon monoxide 26 percent below the goal, and total unburned hydrocarbons 75 percent below the goal. Relative to the JT9D-7 combustor, the oxides of nitrogen were reduced by 58 percent, carbon monoxide emissions were reduced by 69 percent, and total unburned hydrocarbons were reduced by 9 percent. The combustor efficiency and exit temperature profiles were comparable to those of production combustor. Acceleration and starting characteristics were deficient relative to the production engine.

Roberts, R.↗
NASA NTRS1977

An explicit solution to the exoatmospheric powered flight guidance and trajectory optimization problem for rocket propelled vehicles

A derivation of an explicit solution to the two point boundary-value problem of exoatmospheric guidance and trajectory optimization is presented. Fixed initial conditions and continuous burn, multistage thrusting are assumed. Any number of end conditions from one to six (throttling is required in the case of six) can be satisfied in an explicit and practically optimal manner. The explicit equations converge for off nominal conditions such as engine failure, abort, target switch, etc. The self starting, predictor/corrector solution involves no Newton-Rhapson iterations, numerical integration, or first guess values, and converges rapidly if physically possible. A form of this algorithm has been chosen for onboard guidance, as well as real time and preflight ground targeting and trajectory shaping for the NASA Space Shuttle Program.

Jaggers, R. F.↗
NASA NTRS1984

Supersonic-Nozzle Shock-Wave Analysis

Analytical procedure used to modify design of high-pressure-ratio nozzles to reduce vibration during start-up and shutdown. Nozzles used in jet engines, laser nozzles and diffusers, wind tunnels, gas turbines and rocket engines.

Wagner, W. R.↗
NASA NTRS1992

Domain modeling and grid generation for multi-block structured grids with application to aerodynamic and hydrodynamic configurations

About five years ago, a joint development was started of a flow simulation system for engine-airframe integration studies on propeller as well as jet aircraft. The initial system was based on the Euler equations and made operational for industrial aerodynamic design work. The system consists of three major components: a domain modeller, for the graphical interactive subdivision of flow domains into an unstructured collection of blocks; a grid generator, for the graphical interactive computation of structured grids in blocks; and a flow solver, for the computation of flows on multi-block grids. The industrial partners of the collaboration and NLR have demonstrated that the domain modeller, grid generator and flow solver can be applied to simulate Euler flows around complete aircraft, including propulsion system simulation. Extension to Navier-Stokes flows is in progress. Delft Hydraulics has shown that both the domain modeller and grid generator can also be applied successfully for hydrodynamic configurations. An overview is given about the main aspects of both domain modelling and grid generation.

Spekreijse, S. P.↗
NASA NTRS1992

NASA and CFD - Making investments for the future

From a NASA perspective, CFD is a new tool for fluid flow simulation and prediction with virtually none of the inherent limitations of other ground-based simulation techniques. A primary goal of NASA's CFD research program is to develop efficient and accurate computational techniques for utilization in the design and analysis of aerospace vehicles. The program in algorithm development has systematically progressed through the hierarchy of engineering simplifications of the Navier-Stokes equations, starting with the inviscid formulations such as transonic small disturbance, full potential, and Euler.

Hessenius, Kristin A.↗
NASA NTRS1934

A Comparison of Several Methods of Measuring Ignition Lag in a Compression-ignition Engine

The ignition lag of a fuel oil in the combustion chamber of a high speed compression-ignition engine was measured by three different methods. The start of injection of the fuel as observed with a Stoborama was taken as the start of the period of ignition lag in all cases. The end of the period of ignition lag was determined by observation of the appearance of incandescence in the combustion chamber, by inspection of a pressure-time card for evidence of pressure rise, and by analysis of the indicator card for evidence of the combustion of a small but definite quantity of fuel. A comparison of the values for ignition lags obtained by these three methods indicates that the appearance of incandescence is later than other evidences of the start of combustion, that visual inspection of a pressure-time diagram gives consistent and usable values with a minimum requirement of time and/or apparatus, and that analysis of the indicator card is not worth while for ignition lag alone.

Spanogle, J A↗
NASA NTRS2001

Development and Demonstration of a 25 Watt Thermophotovoltaic Power Source for a Hybrid Power System

The development of a propane-fueled, 25 W thermophotovoltaic (TPV) power source for use in a hybrid power system is described. The TPV power source uses a platinum emitting surface with an anti-reflective coating to radiate to gallium antimonide photocells, which converts the radiation to electric power. The development program started with the design and fabrication of an engineering prototype system. This was used as a component development vehicle to develop the technologies for the various components. A 25 W demonstration prototype was then designed and fabricated using the most advanced component approaches. The designs and test results from this development program are discussed.

Doyle, Edward↗
NASA NTRS1998

Using Immersive Virtual Environments for Certification

Immersive virtual environments (VEs) technology has matured to the point where it can be utilized as a scientific and engineering problem solving tool. In particular, VEs are starting to be used to design and evaluate safety-critical systems that involve human operators, such as flight and driving simulators, complex machinery training, and emergency rescue strategies.

virtual reality software certification human-compu↗
NASA NTRS2024

SUPIN: A Tool for the Aerodynamic Design and Analysis of Supersonic Inlets

The Supersonic Inlet Design and Analysis Tool (SUPIN) is a computational program that performs geometric modeling and aerodynamic design and analysis of a set of supersonic inlets whose types include the axisymmetric pitot, three-dimensional pitot, axisymmetric, two-dimensional, two-dimensional bifurcated-duct, and streamline-traced inlets. The inlets are modeled by dividing the inlet flowpath into components that start at the freestream and ends at the engine face. The primary components of the inlet include the external supersonic diffuser, throat section, and subsonic diffuser. Each component is characterized by a set of inputs that include geometric and aerodynamic factors. Examples of the geometric factors are angles, lengths, area ratios, and coordinates that describe the geometry of the components. Examples of the aerodynamic factors are Mach numbers, flow ratios, and pressure ratios that set constraints or desired aerodynamic properties of the component. These input factors are specified within a text-based input data file that is read into SUPIN. SUPIN is a Fortran 95 program that performs the inlet design and analysis as a batch or non-interactive process. SUPIN writes output data files describing the inlet geometry and aerodynamic performance. SUPIN uses analytic, empirical, and numerical methods to perform the inlet design and analysis. The inlet geometry is constructed using planar and three-dimensional geometry constructs and is represented by coordinates, angles, areas, profiles, and surfaces. Surfaces are output in the form of a Plot3D surface grid file. SUPIN does have some capability to automatically generate planar and three-dimensional, multi-block, structured grids for computational fluid dynamics (CFD) analysis. The aerodynamic analysis performed by SUPIN computes quasi-one-dimensional properties, such as Mach numbers, pressures, and temperatures, through the inlet flowpath. Planar aerodynamic properties, such as those for shock and Mach waves, are computed for the supersonic compression of the external supersonic diffuser and throat section. The inlet performance is characterized by the inlet flow rates, total pressure recovery, and inlet drag. The primary intent of this document is to describe the usage of SUPIN. The methods used within SUPIN are documented and may serve as a useful reference for understanding the aerodynamic design and analysis of supersonic inlets.

supersonic inlets↗
NASA NTRS2016

MoonRIDERS: NASA and Hawaiis Innovative Lunar Surface Flight Experiment for Landing in Late 2017

Recently, NASA Kennedy Space Center, Hawaii's state aerospace agency PISCES, and two Hawaii high schools Iolani and Kealakehe have come together in a unique collaboration called MoonRIDERS. This strategic partnership will allow Hawaii students to participate directly in sending a science experiment to the surface of the moon. The MoonRIDERS project started in the spring of 2014, with each institution responsible for its own project costs and activities. PISCES, given its legislative direction in advancing planetary surface systems, saw this collaboration as an important opportunity to inspire a young generation and encourage STEM (Science, Technology, Engineering, and Mathematics) learning. Under the guidance of PISCES and NASA, the students will be involved hands-on from start to finish in the engineering, testing, and validation of a space technology called the Electrodynamic Dust Shield (EDS). Dust is a critical issue for space exploration, as evidenced by the Apollo lunar missions and Mars rovers and landers. Dust creates a number of problems for humans and hardware, including inhalation, mechanical interference, wear and tear on spacesuits, inhibition of heat transfer on radiators, and reduced efficiency of solar panels. To address this, the EDS is designed to work on a variety of materials, and functions by generatingelectrodynamic fields to clear away the dust. The Google Lunar XPRIZE (GLXP), a space competition "designed to inspire pioneers to do robotic space transport on a budget," serves as a likely method for the MoonRIDERS to get their project to the moon. The EDS would potentially be flown as a hosted payload on a competitor's lander (still to be chosen). This briefing will provide an overview of the technology, the unique partnership, progress update and testing leading to this flight opportunity.

Payload experiment↗
NASA NTRS1999

A Synopsis of Ion Propulsion Development Projects in the United States: SERT 1 to Deep Space I

The historical background and characteristics of the experimental flights of ion propulsion systems and the major ground-based technology demonstrations were reviewed. The results of the first successful ion engine flight in 1964, SERT I which demonstrated ion beam neutralization, are discussed along with the extended operation of SERT II starting in 1970. These results together with the technology employed on the early cesium engine flights. the Applications Technology Satellite (ATS) series, and the ground-test demonstrations, have provided the evolutionary path for the development of xenon ion thruster component technologies, control systems, and power circuit implementations. In the 1997-1999 period, the communication satellite flights using ion engine systems and the Deep Space I flight confirmed that these auxiliary and primary propulsion systems have advanced to a high-level of flight-readiness.

Sovey, James S.↗
NASA NTRS1992

Technology test bed engine real-time failure control

The Real-Time Failure Control (RTFC) program involves development of a failure detection algorithm, for the Space Shuttle Main Engine (SSME). This failure detection approach is signal-based and entails monitoring SSME measurement signals based on predetermined as well as on-line computed mean and standard deviation values. Twenty-four engine measurements are monitored in the algorithm and provisions are made to add more parameters if needed. Each of the first values of every measurement signal at the algorithm start is checked against safety limits placed around a pre-computed engine-to-engine mean value (MV) with a bandwidth equal to a given multiple of the pre-computed standard deviation (SD). If several parameters are out of the bounds of these limits a failure is signaled. During the first two seconds (after algorithm start) a moving average (MA) and a SD is computed on-line in real-time. The moving average of each parameter is computed by averaging the incoming signal measurement with the four most recent previous signal measurements. The moving average is updated at every sampling interval (40 msec) and is checked against a similar safety band around the initial signal value for each parameter. If several anomalies are registered, a failure is signaled by the algorithm. At the end of the two-second interval the MA is fixed as the mean value for the rest of the algorithm operation and a safety band is placed above and below this value equal to a multiple of the computed SD. However, the safety band is adjusted by adjusting the mean value when propellant tank repressurization and venting take place. 'Influence Coefficients' are used to make the necessary adjustments to the safety limits of those parameters that are affected by repressurization and venting or valve closure and opening. The MA is, in both cases, continuously updated and checked against the safety band. Once more, if several parameters exceed the limits a failure is signaled. At the start of every scheduled power transient the algorithm is stopped. It is re-initiated after two seconds from the termination of the power transient and the process is repeated. The final report is divided into four major sections. The most encompassing of all is the discussion section that has sub-sections on: (1) RTFC algorithm development, (2) RTFC simulations; (3) RTFC current limitations; and (4) enhancements planned for.

Panossian, Hagop V.↗
NASA NTRS1975

Control logic for a 30 cm diameter ion thruster

The various missions to be performed by the 30-cm diameter mercury bombardment thruster engine are discussed. The operating constraints imposed by the thermal environment, allowable time to reach steady state operation, and the number of start-ups required are examined. The variety of requirements is further analyzed for the impact on the basic control logic for the engine. The control logic is divided into the start-up, run, and shutdown modes of operation. The start-up mode is reported.

Bechtel, R. T.↗
NASA NTRS1971

Status of nuclear light bulb engine design and performance analyses.

The dynamic response is discussed of a nuclear light bulb engine to perturbations at full-power operating conditions, both with and without control systems. Start-up and shutdown procedures are described, along with the provisions for residual afterheat removal. Problems associated with uprated engine designs are examined. The results of some analytical studies indicate that satisfactory control can be obtained by varying the fuel injection rate by means of a control valve which responds to the rate of change of neutron flux, or to the difference between the instantaneous and steady-state neutron fluxes.

Bauer, H. E.↗