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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS1997

A Performance Comparison of Pulsed Plasma Thruster Electrode Configurations

Pulsed plasma thrusters are currently planned on two small satellite missions and proposed for a third. In these missions, the pulsed plasma thruster's unique characteristics will be used variously to provide propulsive attitude control, orbit raising, translation, and precision positioning. Pulsed plasma thrusters are attractive for small satellite applications because they are essentially stand alone devices which eliminate the need for toxic and/or distributed propellant systems. Pulsed plasma thrusters also operate at low power and over a wide power range without loss of performance. As part of the technical development required for the noted missions, an experimental program to optimize performance with respect to electrode configuration was undertaken. One of the planned missions will use pulsed plasma thrusters for orbit raising requiring relatively high thrust and previously tested configurations did not provide this. Also, higher capacitor energies were tested than previously tried for this mission. Multiple configurations were tested and a final configuration was selected for flight hardware development. This paper describes the results of the electrode optimization in detail.

Arrington, Lynn A.↗
NASA NTRS1999

New Approaches to HSCT Multidisciplinary Design and Optimization

New approaches to MDO have been developed and demonstrated during this project on a particularly challenging aeronautics problem- HSCT Aeroelastic Wing Design. To tackle this problem required the integration of resources and collaboration from three Georgia Tech laboratories: ASDL, SDL, and PPRL, along with close coordination and participation from industry. Its success can also be contributed to the close interaction and involvement of fellows from the NASA Multidisciplinary Analysis and Optimization (MAO) program, which was going on in parallel, and provided additional resources to work the very complex, multidisciplinary problem, along with the methods being developed. The development of the Integrated Design Engineering Simulator (IDES) and its initial demonstration is a necessary first step in transitioning the methods and tools developed to larger industrial sized problems of interest. It also provides a framework for the implementation and demonstration of the methodology. Attachment: Appendix A - List of publications. Appendix B - Year 1 report. Appendix C - Year 2 report. Appendix D - Year 3 report. Appendix E - accompanying CDROM.

Schrage, Daniel P.↗
NASA NTRS2000

Parametric Model of an Aerospike Rocket Engine

A suite of computer codes was assembled to simulate the performance of an aerospike engine and to generate the engine input for the Program to Optimize Simulated Trajectories. First an engine simulator module was developed that predicts the aerospike engine performance for a given mixture ratio, power level, thrust vectoring level, and altitude. This module was then used to rapidly generate the aerospike engine performance tables for axial thrust, normal thrust, pitching moment, and specific thrust. Parametric engine geometry was defined for use with the engine simulator module. The parametric model was also integrated into the iSIGHT multidisciplinary framework so that alternate designs could be determined. The computer codes were used to support in-house conceptual studies of reusable launch vehicle designs.

Korte, J. J.↗
NASA NTRS2000

Parametric Model of an Aerospike Rocket Engine

A suite of computer codes was assembled to simulate the performance of an aerospike engine and to generate the engine input for the Program to Optimize Simulated Trajectories. First an engine simulator module was developed that predicts the aerospike engine performance for a given mixture ratio, power level, thrust vectoring level, and altitude. This module was then used to rapidly generate the aerospike engine performance tables for axial thrust, normal thrust, pitching moment, and specific thrust. Parametric engine geometry was defined for use with the engine simulator module. The parametric model was also integrated into the iSIGHTI multidisciplinary framework so that alternate designs could be determined. The computer codes were used to support in-house conceptual studies of reusable launch vehicle designs.

Korte, J. J.↗
NASA NTRS2000

Analytical and Computational Properties of Distributed Approaches to MDO

Historical evolution of engineering disciplines and the complexity of the MDO problem suggest that disciplinary autonomy is a desirable goal in formulating and solving MDO problems. We examine the notion of disciplinary autonomy and discuss the analytical properties of three approaches to formulating and solving MDO problems that achieve varying degrees of autonomy by distributing the problem along disciplinary lines. Two of the approaches-Optimization by Linear Decomposition and Collaborative Optimization-are based on bi-level optimization and reflect what we call a structural perspective. The third approach, Distributed Analysis Optimization, is a single-level approach that arises from what we call an algorithmic perspective. The main conclusion of the paper is that disciplinary autonomy may come at a price: in the bi-level approaches, the system-level constraints introduced to relax the interdisciplinary coupling and enable disciplinary autonomy can cause analytical and computational difficulties for optimization algorithms. The single-level alternative we discuss affords a more limited degree of autonomy than that of the bi-level approaches, but without the computational difficulties of the bi-level methods. Key Words: Autonomy, bi-level optimization, distributed optimization, multidisciplinary optimization, multilevel optimization, nonlinear programming, problem integration, system synthesis

Alexandrov, Natalia M.↗
NASA NTRS1994

Investigation of Methods for the Structural Weight Analysis of a Mach 2.4 Axisymmetric Inlet

Structural design and analysis tools appropriate for estimating the structural weight of an axisymmetric inlet designed for Mach 2.4 cruise were evaluated. Little information regarding the inlet mechanical design is available in the preliminary design phase, so it is necessary to first develop a reasonable structural design before estimating the inlet weight. The Internally Pressurized Structure Synthesis and Optimization (IPSSO) program, employing an analytical approach, was chosen for evaluation due to its combined design and analysis capabilities. The inlet design produced by IPSSO was then analyzed using the NASTRAN finite element program. The finite element analysis was performed to help identify the limitations of the analytically based code as well as to evaluate NASTRAN for this application. Comparison between the IPSSO inlet weight and that of a similar inlet developed by the Boeing Commercial Airplane Group was also made. Program evaluation concluded that the combined use of IPSSO to create an initial design and NASTRAN to perform a numerical analysis would provide the capability to evaluate a limited number of inlet design The development of a new tool for the minimum weight design and analysis of inlet structures would be required for greater flexibility in evaluating inlet conceptual designs.

Nadell, Shari-Beth↗
NASA NTRS2001

Development of a Mars Airplane Entry, Descent, and Flight Trajectory

An entry, descent, and flight (EDF) trajectory profile for a Mars airplane mission is defined as consisting of the following elements: ballistic entry of an aeroshell; supersonic deployment of a decelerator parachute; subsonic release of a heat shield; release, unfolding, and orientation of an airplane to flight attitude; and execution of a pull up maneuver to achieve trimmed, horizontal flight. Using the Program to Optimize Simulated Trajectories (POST) a trajectory optimization problem was formulated. Model data representative of a specific Mars airplane configuration, current models of the Mars surface topography and atmosphere, and current estimates of the interplanetary trajectory, were incorporated into the analysis. The goal is to develop an EDF trajectory to maximize the surface-relative altitude of the airplane at the end of a pull up maneuver, while subject to the mission design constraints. The trajectory performance was evaluated for three potential mission sites and was found to be site-sensitive. The trajectory performance, examined for sensitivity to a number of design and constraint variables, was found to be most sensitive to airplane mass, aerodynamic performance characteristics, and the pull up Mach constraint. Based on the results of this sensitivity study, an airplane-drag optimized trajectory was developed that showed a significant performance improvement.

Murray, James E.↗
NASA NTRS2001

Overview of SBIR Phase II Work on Hollow Graphite Fibers

Ultra-Lightweight materials are enabling for producing space based optical components and support structures. Heretofore, innovative designs using existing materials has been the approach to produce lighter-weight optical systems. Graphite fiber reinforced composites, because of their light weight, have been a material of frequent choice to produce space based optical components. Hollow graphite fibers would be lighter than standard solid graphite fibers and, thus, would save weight in optical components. The Phase I SBIR program demonstrated it is possible to produce hollow carbon fibers that have strengths up to 4.2 GPa which are equivalent to commercial fibers, and composites made from the hollow fibers had substantially equivalent composite strengths as commercial fiber composites at a 46% weight savings. The Phase II SBIR program will optimize processing and properties of the hollow carbon fiber and scale-up processing to produce sufficient fiber for fabricating a large ultra-lightweight mirror for delivery to NASA. Information presented here includes an overview of the strength of some preliminary hollow fibers, photographs of those fibers, and a short discussion of future plans.

Stallcup, Michael↗
NASA NTRS2002

Developments in Hollow Graphite Fiber Technology

Hollow graphite fibers will be lighter than standard solid graphite fibers and, thus, will save weight in optical components. This program will optimize the processing and properties of hollow carbon fibers developed by MER and to scale-up the processing to produce sufficient fiber for fabricating a large ultra-lightweight mirror for delivery to NASA.

Stallcup, Michael↗
NASA NTRS1985

Power Station Design

Kuljian Corporation provides design engineering and construction management services for power generating plants in more than 20 countries. They used WASP (Calculating Water and Steam Properties), a COSMIC program to optimize power station design. This enabled the company to substantially reduce lead time and software cost in a recent design project.

Source record↗
NASA NTRS2002

Mars Smart Lander Simulations for Entry, Descent, and Landing

Two primary simulations have been developed and are being updated for the Mars Smart Lander Entry, Descent, and Landing (EDL). The high fidelity engineering end-to-end EDL simulation that is based on NASA Langley's Program to Optimize Simulated Trajectories (POST) and the end-to-end real-time, hardware-in-the-loop simulation testbed, which is based on NASA JPL's (Jet Propulsion Laboratory) Dynamics Simulator for Entry, Descent and Surface landing (DSENDS). This paper presents the status of these Mars Smart Lander EDL end-to-end simulations at this time. Various models, capabilities, as well as validation and verification for these simulations are discussed.

Striepe, S. A.↗
NASA NTRS1981

Spinoff from a Moon Tool

Portable self-contained drill capable of extracting core samples as much as 10 feet below the surface was needed for the astronauts. Black & Decker used a specially developed computer program to optimize the design of the drill's motor and insure minimal power consumption. Refinement of the original technology led to the development of a cordless miniature vacuum cleaner called the Dustbuster. It has no hose, no cord, is 14 inches long, and also comes with a storage bracket that also serves as a recharger; plugs into a home outlet that charges the nickel cadmium batteries when not in use. Other home use cordless instruments include drills, shrub trimmers and grass shears. Company also manufactures a number of cordless tools used in the sheet metal automobile and construction industries, and a line of cordless orthopedic instruments.

Source record↗
NASA NTRS1998

An Atmospheric Guidance Algorithm Testbed for the Mars Surveyor Program 2001 Orbiter and Lander

An Atmospheric Flight Team was formed by the Mars Surveyor Program '01 mission office to develop aerocapture and precision landing testbed simulations and candidate guidance algorithms. Three- and six-degree-of-freedom Mars atmospheric flight simulations have been developed for testing, evaluation, and analysis of candidate guidance algorithms for the Mars Surveyor Program 2001 Orbiter and Lander. These simulations are built around the Program to Optimize Simulated Trajectories. Subroutines were supplied by Atmospheric Flight Team members for modeling the Mars atmosphere, spacecraft control system, aeroshell aerodynamic characteristics, and other Mars 2001 mission specific models. This paper describes these models and their perturbations applied during Monte Carlo analyses to develop, test, and characterize candidate guidance algorithms.

Striepe, Scott A.↗
NASA NTRS2004

Aerocapture Performance Analysis for a Neptune-Triton Exploration Mission

A systems analysis has been conducted for a Neptune-Triton Exploration Mission in which aerocapture is used to capture a spacecraft at Neptune. Aerocapture uses aerodynamic drag instead of propulsion to decelerate from the interplanetary approach trajectory to a captured orbit during a single pass through the atmosphere. After capture, propulsion is used to move the spacecraft from the initial captured orbit to the desired science orbit. A preliminary assessment identified that a spacecraft with a lift to drag ratio of 0.8 was required for aerocapture. Performance analyses of the 0.8 L/D vehicle were performed using a high fidelity flight simulation within a Monte Carlo executive to determine mission success statistics. The simulation was the Program to Optimize Simulated Trajectories (POST) modified to include Neptune specific atmospheric and planet models, spacecraft aerodynamic characteristics, and interplanetary trajectory models. To these were added autonomous guidance and pseudo flight controller models. The Monte Carlo analyses incorporated approach trajectory delivery errors, aerodynamic characteristics uncertainties, and atmospheric density variations. Monte Carlo analyses were performed for a reference set of uncertainties and sets of uncertainties modified to produce increased and reduced atmospheric variability. For the reference uncertainties, the 0.8 L/D flatbottom ellipsled vehicle achieves 100% successful capture and has a 99.87 probability of attaining the science orbit with a 360 m/s V budget for apoapsis and periapsis adjustment. Monte Carlo analyses were also performed for a guidance system that modulates both bank angle and angle of attack with the reference set of uncertainties. An alpha and bank modulation guidance system reduces the 99.87 percentile DELTA V 173 m/s (48%) to 187 m/s for the reference set of uncertainties.

Starr, Brett R.↗
NASA NTRS2005

Aerocapture Performance Analysis of A Venus Exploration Mission

A performance analysis of a Discovery Class Venus Exploration Mission in which aerocapture is used to capture a spacecraft into a 300km polar orbit for a two year science mission has been conducted to quantify its performance. A preliminary performance assessment determined that a high heritage 70 sphere-cone rigid aeroshell with a 0.25 lift to drag ratio has adequate control authority to provide an entry flight path angle corridor large enough for the mission s aerocapture maneuver. A 114 kilograms per square meter ballistic coefficient reference vehicle was developed from the science requirements and the preliminary assessment s heating indicators and deceleration loads. Performance analyses were conducted for the reference vehicle and for sensitivity studies on vehicle ballistic coefficient and maximum bank rate. The performance analyses used a high fidelity flight simulation within a Monte Carlo executive to define the aerocapture heating environment and deceleration loads and to determine mission success statistics. The simulation utilized the Program to Optimize Simulated Trajectories (POST) that was modified to include Venus specific atmospheric and planet models, aerodynamic characteristics, and interplanetary trajectory models. In addition to Venus specific models, an autonomous guidance system, HYPAS, and a pseudo flight controller were incorporated in the simulation. The Monte Carlo analyses incorporated a reference set of approach trajectory delivery errors, aerodynamic uncertainties, and atmospheric density variations. The reference performance analysis determined the reference vehicle achieves 100% successful capture and has a 99.87% probability of attaining the science orbit with a 90 meters per second delta V budget for post aerocapture orbital adjustments. A ballistic coefficient trade study conducted with reference uncertainties determined that the 0.25 L/D vehicle can achieve 100% successful capture with a ballistic coefficient of 228 kilograms per square meter and that the increased ballistic coefficient increases post aerocapture V budget to 134 meters per second for a 99.87% probability of attaining the science orbit. A trade study on vehicle bank rate determined that the 0.25 L/D vehicle can achieve 100% successful capture when the maximum bank rate is decreased from 30 deg/s to 20 deg/s. The decreased bank rate increases post aerocapture delta V budget to 102 meters per second for a 99.87% probability of attaining the science orbit.

Starr, Brett R.↗
NASA NTRS1994

A Method For Modeling Discontinuities In A Microwave Coaxial Transmission Line

A methodology for modeling discountinuities in a coaxial transmission line is presented. The method uses a none-linear least squares fit program to optimize the fit between a theoretical model and experimental data. When the method was applied for modeling discontinuites in a damaged S-band antenna cable, excellent agreement was obtained.

Galileo S-band antenna discontinuities↗
NASA NTRS2007

POST2 End-To-End Descent and Landing Simulation for the Autonomous Landing and Hazard Avoidance Technology Project

The Program to Optimize Simulated Trajectories II (POST2) is used as a basis for an end-to-end descent and landing trajectory simulation that is essential in determining the design and performance capability of lunar descent and landing system models and lunar environment models for the Autonomous Landing and Hazard Avoidance Technology (ALHAT) project. This POST2-based ALHAT simulation provides descent and landing simulation capability by integrating lunar environment and lander system models (including terrain, sensor, guidance, navigation, and control models), along with the data necessary to design and operate a landing system for robotic, human, and cargo lunar-landing success. This paper presents the current and planned development and model validation of the POST2-based end-to-end trajectory simulation used for the testing, performance and evaluation of ALHAT project system and models.

Fisher, Jody l.↗
NASA NTRS2008

Ares-I-X Vehicle Preliminary Range Safety Malfunction Turn Analysis

Ares-I-X is the designation given to the flight test version of the Ares-I rocket (also known as the Crew Launch Vehicle - CLV) being developed by NASA. As part of the preliminary flight plan approval process for the test vehicle, a range safety malfunction turn analysis was performed to support the launch area risk assessment and vehicle destruct criteria development processes. Several vehicle failure scenarios were identified which could cause the vehicle trajectory to deviate from its normal flight path, and the effects of these failures were evaluated with an Ares-I-X 6 degrees-of-freedom (6-DOF) digital simulation, using the Program to Optimize Simulated Trajectories Version 2 (POST2) simulation framework. The Ares-I-X simulation analysis provides output files containing vehicle state information, which are used by other risk assessment and vehicle debris trajectory simulation tools to determine the risk to personnel and facilities in the vicinity of the launch area at Kennedy Space Center (KSC), and to develop the vehicle destruct criteria used by the flight test range safety officer. The simulation analysis approach used for this study is described, including descriptions of the failure modes which were considered and the underlying assumptions and ground rules of the study, and preliminary results are presented, determined by analysis of the trajectory deviation of the failure cases, compared with the expected vehicle trajectory.

Beaty, James R.↗