SEARCH · Search NASA
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.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
Contents include the following: PCB design with HDL designer, design process and schematic capture - symbols and diagrams: 1. Motivation: time savings, money savings, simplicity. 2. Approach: use single tool PCB for FPGA design, more FPGA designs than PCB designers. 3. Use HDL designer for schematic capture.
Mentor's DxDesigner and Expedition schematic capture and printed circuit board tools were chosen to implement a custom high speed signal processing board containing many high pin count Field Programmable Gate Arrays and many high speed serial connections with data rates over 2 Gigasamples/sec. The methodology used to place the parts and route the board involved the interaction of both the DxDesigner and Expedition tools. The basic design philosophy was to specify as much as possible through design constraints at the schematic level. This paper will explore implementing that philosophy in both tools to facilitate part placement and trace routing.
An investigation of the aeroheating environment of the Project Orion Crew Entry Vehicle has been performed in the Arnold Engineering Development Center Tunnel 9. Data were measured on a approx. 3.5% scale model (0.1778m/7-inch diam.) of the vehicle using coaxial thermocouples in the Mach 8 and Mach 10 nozzles of Tunnel 9. Runs were performed at free stream Reynolds numbers of 1 106/ft to 20 10(exp 6)/ft in the Mach 10 nozzle and 8 10(exp 6)/ft to 48 10(exp 6)/ft in the Mach 8 nozzle. The test gas in Tunnel 9 is pure N2, which at these operating conditions remains un-dissociated and may be treated as a perfect gas. At these conditions, laminar, transitional, and turbulent flow was produced on the model at Mach 10, and transitional and turbulent conditions were produced on the model at Mach 8. The majority of runs were made on a clean, smooth-surface model configuration and a limited number of runs were made in which inserts with varying boundary-layer trips configurations were used to force the occurrence of transition. Laminar and turbulent predictions were generated for all wind tunnel test conditions and comparisons were performed with the data for the purpose of helping to define uncertainty margins for the computational method. Data from both the wind tunnel test and the computations are presented herein. Figure 1 shows a schematic of the thermocouple locations on the model and figures 2 and 3 show a photo and schematic of the AEDC Hypervelocity Tunnel 9. Figure 4 shows a typical grid used in the computations. From the comparisons shown in figures 5 through 8 it was concluded that for perfect-gas conditions, the computations could predict either fully-laminar or full-turbulent flow to within +/-10% of the experimental data. The experimental data showed that transition began on the leeside of the heatshield at a free stream Reynolds number of 9 10(exp 6)/ft in the Mach 10 nozzle and fully-developed turbulent flow was produced at 20 10(exp 6)/ft. In the Mach 8 nozzle, transition on the leeside of the heat-shield was observed for all test conditions, and full-developed turbulent flow occurred at a free stream Reynolds number of 18 10(exp 6)/ft. On the aftbody of the vehicle no evidence of turbulence was detected at Mach 10 conditions, and at Mach 8 conditions, transition appeared to begin on the windside of the aftbody at free stream Reynolds number of 18x10(exp 6)/ft with fully-developed turbulent flow occurring only at the highest test condition of 48x10(exp 6)/ft.
As part of NASA s initiative to develop an advanced portable life support system (PLSS), a baseline schematic has been chosen that includes gaseous oxygen in a closed circuit ventilation configuration. Supply oxygen enters the suit at the back of the helmet and return gases pass over the astronaut s body to be extracted at the astronaut s wrists and ankles through the liquid cooling and ventilation garment (LCVG). The extracted gases are then treated using a rapid cycling amine (RCA) system for carbon dioxide and water removal and activated carbon for trace gas removal before being mixed with makeup oxygen and reintroduced into the helmet. Thermal control is provided by a suit water membrane evaporator (SWME). As an extension of the original schematic development, NASA evaluated several Helmet Exhalation Capture System (HECS) configurations as alternatives to the baseline. The HECS configurations incorporate the use of full contact masks or non-contact masks to reduce flow requirements within the PLSS ventilation subsystem. The primary scope of this study was to compare the alternatives based on mass and volume considerations; however other design issues were also briefly investigated. This paper summarizes the results of this sizing analysis task.
Preliminary design trades are presented for liquid hydrogen fuel systems for remotely-operated, high-altitude aircraft that accommodate three different propulsion options: internal combustion engines, and electric motors powered by either polymer electrolyte membrane fuel cells or solid oxide fuel cells. Mission goal is sustained cruise at 60,000 ft altitude, with duration-aloft a key parameter. The subject aircraft specifies an engine power of 143 to 148 hp, gross liftoff weight of 9270 to 9450 lb, payload of 440 lb, and a hydrogen fuel capacity of 2650 to 2755 lb stored in two spherical tanks (8.5 ft inside diameter), each with a dry mass goal of 316 lb. Hydrogen schematics for all three propulsion options are provided. Each employs vacuum-jacketed tanks with multilayer insulation, augmented with a helium pressurant system, and using electric motor driven hydrogen pumps. The most significant schematic differences involve the heat exchangers and hydrogen reclamation equipment. Heat balances indicate that mission durations of 10 to 16 days appear achievable. The dry mass for the hydrogen system is estimated to be 1900 lb, including 645 lb for each tank. This tank mass is roughly twice that of the advanced tanks assumed in the initial conceptual vehicle. Control strategies are not addressed, nor are procedures for filling and draining the tanks.
Excavating granular materials beneath a vertical jet of gas involves several physical mechanisms. These occur, for example, beneath the exhaust plume of a rocket landing on the soil of the Moon or Mars. We performed a series of experiments and simulations (Figure 1) to provide a detailed view of the complex gas-soil interactions. Measurements taken from the Apollo lunar landing videos (Figure 2) and from photographs of the resulting terrain helped demonstrate how the interactions extrapolate into the lunar environment. It is important to understand these processes at a fundamental level to support the ongoing design of higher fidelity numerical simulations and larger-scale experiments. These are needed to enable future lunar exploration wherein multiple hardware assets will be placed on the Moon within short distances of one another. The high-velocity spray of soil from the landing spacecraft must be accurately predicted and controlled or it could erode the surfaces of nearby hardware. This analysis indicated that the lunar dust is ejected at an angle of less than 3 degrees above the surface, the results of which can be mitigated by a modest berm of lunar soil. These results assume that future lunar landers will use a single engine. The analysis would need to be adjusted for a multiengine lander. Figure 3 is a detailed schematic of the Lunar Module camera calibration math model. In this chart, formulas relating the known quantities, such as sun angle and Lunar Module dimensions, to the unknown quantities are depicted. The camera angle PSI is determined by measurement of the imaged aspect ratio of a crater, where the crater is assumed to be circular. The final solution is the determination of the camera calibration factor, alpha. Figure 4 is a detailed schematic of the dust angle math model, which again relates known to unknown parameters. The known parameters now include the camera calibration factor and Lunar Module dimensions. The final computation is the ejected dust angle, as a function of Lunar Module altitude.
A detailed description of the Nuclear Thermal Rocket Element Environmental Simulator (NTREES) is presented. The contents include: 1) Design Requirements; 2) NTREES Layout; 3) Data Acquisition and Control System Schematics; 4) NTREES System Schematic; and 5) NTREES Setup.
Controlled Technology: 1) Design: preliminary or critical design data, schematics, technical flow charts, SNV code/diagnostics, logic flow diagrams, wirelist, ICDs, detailed specifications or requirements. 2) Development: constraints, computations, configurations, technical analyses, acceptance criteria, anomaly resolution, detailed test plans, detailed technical proposals. 3) Production: process or how-to: assemble, operated, repair, maintain, modify. 4) Manufacturing: technical instructions, specific parts, specific materials, specific qualities, specific processes, specific flow. 5) Operations: how-to operate, contingency or standard operating plans, Ops handbooks. 6) Repair: repair instructions, troubleshooting schemes, detailed schematics. 7) Test: specific procedures, data, analysis, detailed test plan and retest plans, detailed anomaly resolutions, detailed failure causes and corrective actions, troubleshooting, trended test data, flight readiness data. 8) Maintenance: maintenance schedules and plans, methods for regular upkeep, overhaul instructions. 9) Modification: modification instructions, upgrades kit parts, including software
The Sabatier Carbon Dioxide Reduction System (CRA) offers water recovery on a long duration space mission to reduce water resupply. Currently, NASA Johnson Space Center (JSC), NASA Marshall Space Flight Center (MSFC), Hamilton Sundstrand Space Systems International, Inc. (HSSSI), and Southwest Research Institute (SWRI) are working together to develop a Sabatier CRA for the International Space Station (ISS). This effort is being funded by the Office of Biological and Physical Research (Code U)/Advanced Life Support program which is administered by NASA JSC. The Sabatier CRA is the next step in closing the oxygen life support loop on future space missions. The Sabatier reaction combines the waste carbon dioxide (recovered from crew metabolism) with waste hydrogen (a byproduct of electrolysis to produce oxygen) to produce water and methane (CH4). On ISS, the methane would be vented overboard, however the methane can be utilized for propulsion during a planetary exploration mission. Based on a crew size of 7-equivalent people, the Sabatier CRA can produce as much a 2000 lb/year water. Use of the Sabatier CRA will significantly reduce the amount of water that needs to be resupplied to the ISS on a yearly basis, at a tremendous cost saving to the program. Additionally, by recycling this additional water, the Sabatier CRA enables additional launch capacity for science experiments to be brought up to the ISS. The NASA/Industry team noted above has been working to reduce technical risks associated with the Sabatier CRA system. To date the technical risks have been considerably reduced, bringing the Technology Readiness Level (TRL) from TRL 4 to TRL 5/6. In doing so, the team has developed the system schematic, system models, control scheme, produced engineering development unit (EDU) hardware, performed limited integration testing of the EDU's and verified system modeling through testing. Additionally, the system schematic has been evaluated for failure modes and hazards and had a successful technical review by the NASA Safety Board. The current focus is now related to development of the water/methane phase separator, liquid sensor and CO2 compressor piston seal life. The overall goal of the current effort is to bring the system up to a TRL6 by the end of GFY04. Although the Sabatier CRA is not currently baselined for use on the ISS, its benefits are significant enough such that volume within the Oxygen Generation System rack has been reserved for future installation. The value of the water the CRA recover will allow NASA the additional crew time and payload needed to pursue its mission of scientific research.
During my internship at NASA, I was a model developer for Ground Support Equipment (GSE). The purpose of a model developer is to develop and unit test model component libraries (fluid, electrical, gas, etc.). The models are designed to simulate software for GSE (Ground Special Power, Crew Access Arm, Cryo, Fire and Leak Detection System, Environmental Control System (ECS), etc. ~.) before they are implemented into hardware. These models support verifying local control and remote software for End-Item Software Under Test (SUT). The model simulates the physical behavior (function, state, limits and 110) of each end-item and it's dependencies as defined in the Subsystem Interface Table, Software Requirements & Design Specification (SRDS), Ground Integrated Schematic (GIS), and System Mechanical Schematic.(SMS). The software of each specific model component is simulated through MATLAB's Simulink program. The intensiv~ model development life cycle is a.s follows: Identify source documents; identify model scope; update schedule; preliminary design review; develop model requirements; update model.. scope; update schedule; detailed design review; create/modify library component; implement library components reference; implement subsystem components; develop a test script; run the test script; develop users guide; send model out for peer review; the model is sent out for verific~tionlvalidation; if there is empirical data, a validation data package is generated; if there is not empirical data, a verification package is generated; the test results are then reviewed; and finally, the user. requests accreditation, and a statement of accreditation is prepared. Once each component model is reviewed and approved, they are intertwined together into one integrated model. This integrated model is then tested itself, through a test script and autotest, so that it can be concluded that all models work conjointly, for a single purpose. The component I was assigned, specifically, was a fluid component, a discrete pressure switch. The switch takes a fluid pressure input, and if the pressure is greater than a designated cutoff pressure, the switch would stop fluid flow.
This project involves creating software for support equipment used on the Space l aunch System (SLS). The goal is to create applications and displays that will be used to remotely operate equipment from the firing room and will continue to support the SLS launch vehicle to the extent of its program. These displays include design practices that help to convey information effectively, such as minimizing distractions at normal operating state and displaying intentional distractions during a warning or alarm state. The general practice for creating an operator display is to reduce the detail of unimportant aspects of the display and promote focus on data and dynamic information. These practices include using minimalist design, using muted tones for background colors, using a standard font at a readable text size, displaying alarms visible for Immediate attention, grouping data logically, and displaying data appropriately varying on the type of data. Users of these displays are more likely to stay focused on operating for longer periods by using design practices that reduce eye strain and fatigue. Effective operator displays will improve safety by reducing human errors during operation, which will help prevent catastrophic accidents. This report entails the details of my work on developing remote displays for the Hypergolics ground system. Before developing a prototype display, the design and requirements of the system are outlined and compiled into a document. Then each subsystem has schematic representations drawn tha.t meet the specifications detailed in the document. The schematics are then used as the outline to create display representations of each subsystem. Each display is first tested individually. Then the displays are integrated with a prototype of the master system, and they are tested in a simulated environment then retested in the real environment. Extensive testing is important to ensure the displays function reliably as intended.
This project involves creating software for support equipment used on the Space Launch System (SLS). The goal is to create applications and displays that will be used to remotely operate equipment from the firing room and will continue to support the SLS launch vehicle to the extent of its program. These displays include design practices that help to convey information effectively, such as minimizing distractions at normal operating state and displaying intentional distractions during a warning or alarm state. The general practice for creating an operator display is to reduce the detail of unimportant aspects of the display and promote focus on data and dynamic information. These practices include using minimalist design, using muted tones for background colors, using a standard font at a readable text size, displaying alarms visible for immediate attention, grouping data logically, and displaying data appropriately varying on the type of data. Users of these displays are more likely to stay focused on operating for longer periods by using design practices that reduce eye strain and fatigue. Effective operator displays will improve safety by reducing human errors during operation, which will help prevent catastrophic accidents. This report entails the details of my work on developing remote displays for the Hypergolic fuel servicing system. Before developing a prototype display, the design and requirements of the system are outlined and compiled into a document. Then each subsystem has schematic representations drawn that meet the specifications detailed in the document. The schematics are then used as the outline to create display representations of each subsystem. Each display is first tested individually. Then the displays are integrated with a prototype of the master system, and they are tested in a simulated environment then retested in the real environment. Extensive testing is important to ensure the displays function reliably as intended.
A one-directional dynamic model of a Stewart Platform was developed to assist NASA in analyzing the dynamic response in spacecraft docking operations. A simplified mechanical drawing was created, capturing the physical structure's main features. A simplified schematic diagram was developed in a lumped mass model from the mechanical drawing. Three differential equations were derived according to the schematic diagram. A Simulink diagram was created using MATLAB to represent the three equations. System parameters, including spring constants and masses, are derived in detail from the physical system. The model can be used for further analysis via computer simulation in predicting dynamic response in its main docking direction, i.e., up-and-down motion.
Schematics of solid state sequential camera trigger circuits
Schematic and flow diagrams showing uplink and downlink of Apollo spacecraft automatic checkout equipment
Automated drafting system produces schematic and block diagrams from the design engineers freehand sketches. This system codes conventional drafting symbols and their coordinate locations on standard size drawings for entry on tapes that are used to drive a high speed photocomposition machine.
Schematic diagrams and monitoring information for behavior of silicon integrated microcircuits in electron radiation environment
Single-ended step-up regulator-chopper power supply /employing conventional chopper circuitry/ combines the advantages of the chopper and switching regulator circuits. Schematic of the power supply incorporating the step-up regulator is shown.