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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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EASY5 modeling of a complete spacecraft power processing system is presented. Component models are developed, and several system models including a solar array switching system, a partially-shunted solar array system and COBE system are simulated. The power system's modes of operation, such as shunt mode, battery-charge mode, and battery-discharge mode, are simulated for a complete orbit cycle.
Studies were conducted to determine the performance requirements for projected state-of-the-art SEP spacecrafts boosted by the Shuttle/IUS to perform a rendezvous with the comet Halley and a rendezvous with the comet Encke during its 1977 apparition. The spacecraft model of the standard HILTOP computer program was assumed. Numerical and graphical results summarizing the studies are presented.
The Communication Systems Simulation Laboratory (CSSL) at the National Aeronautics and Space Administration (NASA)/Johnson Space Center (JSC) is tasked to perform spacecraft and ground network communication system simulations. The CSSL has developed simulation tools that model spacecraft communication systems and the space/ground environment in which they operate. This paper is to analyze a spacecraft's very high frequency (VHF) radio signal propagation and the impact to performance when landing in an ocean. Very little research work has been done for VHF radio systems in a maritime environment. Rigorous Radio Frequency (RF) modeling/simulation techniques were employed for various environmental effects. The simulation results illustrate the significance of the environmental effects on the VHF radio system performance.
The Human Interface Branch - EV3 - is evaluating Organic lighting-emitting diodes (OLEDs) as an upgrade for current displays on future spacecraft. OLEDs have many advantages over current displays. Conventional displays require constant backlighting which draws a lot of power, but with OLEDs they generate light themselves. OLEDs are lighter, and weight is always a concern with space launches. OLEDs also grant greater viewing angles. OLEDs have been in the commercial market for almost ten years now. What is not known is how they will perform in a space-like environment; specifically deep space far away from the Earth's magnetosphere. In this environment, the OLEDs can be expected to experience vacuum and galactic radiation. The intern's responsibility has been to prepare the OLED for a battery of tests. Unfortunately, it will not be ready for testing at the end of the internship. That being said much progress has been made: a) Developed procedures to safely disassemble the tablet. b) Inventoried and identified critical electronic components. c) 3D printed a testing apparatus. d) Wrote software in Python that will test the OLED screen while being radiated. e) Built circuits to restart the tablet and the test pattern, and ensure it doesn't fall asleep during radiation testing. f) Built enclosure that will house all of the electronics Also, the intern has been working on a way to take messages from a simulated Caution and Warnings system, process said messages into packets, send audio packets to a multicast address that audio boxes are listening to, and output spoken audio. Currently, Cautions and Warnings use a tone to alert crew members of a situation, and then crew members have to read through their checklists to determine what the tone means. In urgent situations, EV3 wants to deliver concise and specific alerts to the crew to facilitate any mitigation efforts on their part. Significant progress was made on this project: a) Open channel with the simulated Caution and Warning system to acquire messages. b) Configure audio boxes. c) Grab pre-recorded audio files. d) Packetize the audio stream. A third project that was assigned to implement LED indicator modules for an Omnibus project. The Omnibus project is investigating better ways designing lighting for the interior of spacecraft-both spacecraft lighting and avionics box status lighting indication. The current scheme contains too much of the blue light spectrum that disrupts the sleep cycle. The LED indicator modules are to simulate the indicators running on a spacecraft. Lighting data will be gathered by human factors personal and use in a model underdevelopment to model spacecraft lighting. Significant progress was made on this project: Designed circuit layout a) Tested LEDs at LETF. b) Created GUI for the indicators. c) Created code for the Arduino to run that will illuminate the indicator modules.
An interactive computer aided design program used to perform systems level design and analysis of large spacecraft concepts is presented. Emphasis is on rapid design, analysis of integrated spacecraft, and automatic spacecraft modeling for lattice structures. Capabilities and performance of multidiscipline applications modules, the executive and data management software, and graphics display features are reviewed. A single user at an interactive terminal create, design, analyze, and conduct parametric studies of Earth orbiting spacecraft with relative ease. Data generated in the design, analysis, and performance evaluation of an Earth-orbiting large diameter antenna satellite are used to illustrate current capabilities. Computer run time statistics for the individual modules quantify the speed at which modeling, analysis, and design evaluation of integrated spacecraft concepts is accomplished in a user interactive computing environment.
The NASA Orbital Debris Program Office (ODPO) relies on ground-based radar measurements from both the Haystack Ultrawideband Satellite Imaging Radar (HUSIR) and the Goldstone Solar System Radar (Goldstone) to characterize mm to cm debris population in low Earth orbit (LEO). Radar measurements help characterize the size of orbital debris objects, particularly fragmentation debris. However, debris size is not directly measured by radar but inferred from the measured radar cross section (RCS) which depends on several parameters in addition to physical size including electrical conductivity and polarization. To interpret the observed RCS of orbital debris objects detected by radar measurements as physical sizes, NASA uses an empirical size estimation model (SEM) based on laboratory RCS measurements of breakup fragments generated during hypervelocity impact tests as well as some pieces of non-impact-generated “artificial” debris-like objects expected to be representative of the debris population. Since the development of the ODPO SEM, many new materials have been introduced to spacecraft construction. Consequently, ODPO plans to update the radar SEM based on planned laboratory RCS measurements of debris fragments from DebriSat, a ground-based hypervelocity impact experiment conducted in 2014 that consisted of a high-fidelity spacecraft model characteristic of a modern LEO spacecraft. Prior to measuring DebriSat fragments, a set of calibration targets with well-defined geometries and material compositions were measured at The Ohio State University’s ElectroScience Laboratory (OSU-ESL) compact radar range. These calibration measurements help to validate, and understand any limitations of, laboratory measurements of RCS. Calibration targets include idealizations of typical shape categories seen in DebriSat fragments such as nuggets, flat plates, and cylinders. As with DebriSat, calibration target materials were chosen to represent typical modern-day spacecraft components and include stainless steel, aluminum, printed circuit board (PCB) substrate, and carbon fiber-reinforced polymer (CFRP). These materials also represent a wide range of electrical conductivities, which strongly influences measured RCS and inferred target size. The RCS calibration measurements were collected over a frequency sweep from 2 to 18 GHz and stepping through different azimuth angles from 0 to 360 degrees at an elevation of 0 degrees. A second set of calibration measurements is in work consisting of more complex shapes such as bent rods and plates as well as different mounting options including epoxy and a 3D printed holder. These further measurements along with our initial calibration set will inform selection of representative DebriSat fragments for laboratory RCS measurements that will contribute to the planned update to the ODPO radar SEM. An appropriate subset of both the radar calibration and DebriSat samples will also be measured in the ODPO Optical Measurements Center to cross-calibrate size estimates over these different wavelength regimes.
An improved specification of the plasma environment has been developed for use in modeling spacecraft charging. It was developed by statistically analyzing a large part of the LANL Magnetospheric Plasma Analyzer (MPA) data set for ion and electron spectral signature correlation with spacecraft charging, including anisotropies. The objective is to identify a relatively simple characterization of the full particle distributions that yield an accurate predication of the observed charging under a wide variety of conditions.
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.
Spacecraft structural designs are typically verified through a coupled loads analysis (CLA) process, which couples the spacecraft model with the launch vehicle (LV) model to predict low-frequency quasi-static and dynamic responses. The CLA calculations are typically the responsibility of the LV organization, but the spacecraft organization has a vested interest in being able to calculate approximate CLA results during the design of the spacecraft. Because of this, there has long been interest in a method that would allow a spacecraft organization to perform a CLA without access to the full set of LV models and forcing functions. One such method is the Norton-Thevenin Receptance Coupling (NTRC) approach, which is specifically designed to accurately transform LV free accelerations (no payload) into coupled system accelerations (LV plus payload). The purpose of this report is to provide historical context for the NTRC method and compare it with methods that have been used in the past. In particular, it is compared to a frequency-domain substitution method that had been used for a long period of time at the Jet Propulsion Laboratory, and a component-mode-based equivalent to that method.
The Communication Systems Simulation Laboratory (CSSL) at the NASA Johnson Space Center is tasked to perform spacecraft and ground network communication system simulations, design validation, and performance verification. The CSSL has developed simulation tools that model spacecraft communication systems and the space and ground environment in which the tools operate. In this paper, a spacecraft communication system with multiple arrays is simulated. Multiple array combined technique is used to increase the radio frequency coverage and data rate performance. The technique is to achieve phase coherence among the phased arrays to combine the signals at the targeting receiver constructively. There are many technical challenges in spacecraft integration with a high transmit power communication system. The array combining technique can improve the communication system data rate and coverage performances without increasing the system transmit power requirements. Example simulation results indicate significant performance improvement can be achieved with phase coherence implementation.
Thermal scale modeling techniques for Voyager- type spacecraft
A number of phenomena observed on the first orbiting Meteosat satellite attributed to spacecraft charging effects are considered. Design analysis, correlation of anomalies with space environmental data, on ground tests with an engineering model spacecraft, tests on the validity of improvements, and installation of suitable monitors for the second improved flight satellite are discussed.
Material removed from the spacecraft and returned to earth remained on the lunar surface for 31.9 lunations. None of the returned parts received the maximum 10,686 hours of exposure because of shadowing by the planar array antenna, solar panel, thermal control compartments, or other parts of the spacecraft. To determine the actual exposure of specific parts to sunlight, six series of photographs were taken. A one-fifth-scale model spacecraft was oriented to a collimated light source simulating the orientation of Surveyor with the sun. Three cameras were set up to view different parts of the spacecraft. The data obtained from these photographs permitted an evaluation of the effects of exposure to solar radiation on the camera and its parts, the surface sampling scoop, and the strut from the radar altimeter and Doppler velocity sensor.
This paper describes the Dynamics Algorithms for Real-Time Simulation (DARTS) real-time hardware-in-the-loop dynamics simulator for the National Aeronautics and Space Administration's Cassini spacecraft. The spacecraft model consists of a central flexible body with a number of articulated rigid-body appendages. The demanding performance requirements from the spacecraft control system require the use of a high fidelity simulator for control system design and testing. The DARTS algorithm provides a new algorithmic and hardware approach to the solution of this hardware-in-the-loop simulation problem. It is based upon the efficient spatial algebra dynamics for flexible multibody systems. A parallel and vectorized version of this algorithm is implemented on a low-cost, multiprocessor computer to meet the simulation timing requirements.
A mathematical synthesization approach for rapidly modeling and analyzing complex structures comprised of hundreds or thousands of repeating individual structural members, interconnecting hardware, and other components was described. The techniques are applicable to any structure with repeating structural members. The capabilities are embodied within the Interactive Design and Evaluation of Advanced Spacecraft (IDEAS) computer-aided design system. IDEAS has a full range of integrated spacecraft modeling, design, analysis, performance, and cost estimating capabilities which include about 40 technical program and executive, data base, and file management systems. With IDEAS, a single user at an interactive terminal can create, design, analyze, and conduct parametric studies of Earth orbiting spacecraft in a timely, cost-efficient manner. The system is particularly useful in the conceptual design phases of advanced space missions when a multiplicity of concepts must be evaluated. The IDEAS interactive finite-element modeling capabilities are discussed. The mathematical synthesization programs were structured not only to provide a detailed finite-element model of the structure but also to automatically perform the many auxiliary computations needed to analyze the spacecraft in its orbital operational environment.
Static attitude estimation and dynamic attitude estimation are used to describe a gradiometer composed of a number of accelerometers that are used to measure a combination of the local gravity gradient and instrument rotation effects. After a series of measures to isolate the gradient, a global mesh of measurements can be obtained that determine the planetary external gravity potential. Orbital and spacecraft models are developed to determine if, when the gravity potential is known, the same measurements, unsupported by any other information can be used to infer the spacecraft attitude.
The standard AP8 and AE8 models for predicting trapped proton and electron environments have been compared with several sets of flight data to evaluate model uncertainties. Model comparisons are made with flux, dose, and activation measurements made on various U.S. low-Earth orbit satellites (APEX, CRRES, DMSP. LDEF, NOAA) and Space Shuttle flights, on Russian satellites (Photon-8, Cosmos-1887, Cosmos-2044), and on the Russian Mir space station. This report gives a summary of the model-data given in a companion report. Results from the model comparisons with flight data show, for example, that the AP8 model underpredicts the trapped proton flux at low altitudes by a factor of about two (independent of proton energy and solar cycle conditions), and that the AE8 model overpredict the flux in the outer electron belt be an order of magnitude or more.
The standard AP8 and AE8 models for predicting trapped proton and electron environments have been compared with several sets of flight data to evaluate model uncertainties. Model comparisons are made with flux, dose, and activation measurements made on various U.S. low-Earth orbit satellites (APEX, CRRES, DMSP, LDEF, NOAA) and Space Shuttle flights, on Russian satellites (Photon-8, Cosmos-1887, Cosmos-2044), and on the Russian Mir Space Station. This report gives a summary of the model-data comparisons-detailed results are given in a companion report. Results from the model comparisons with flic,ht data show, for example, the AP8 model underpredicts the trapped proton flux at low altitudes by a factor of about two (independent of proton energy and solar cycle conditions), and that the AE8 model overpredicts the flux in the outer electron belt by an order of magnitude or more.