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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 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

An overview of the Solar, Anomalous, and Magnetospheric Particle Explorer (SAMPEX) mission

The scientific objective of the NASA Small-class Explorer Mission SAMPEX is summarized. A brief history of the Small Explorer program is provided along with a description of the SAMPEX project development and structure. The spacecraft and scientific instrument configuration is presented. The orbit of SAMPEX has an altitude of 520 by 670 km and an 82 deg inclination. Maximum possible power is provided by articulated solar arrays that point continuously toward the sun. Highly sensitive science instruments point generally toward the local zenith, especially over the terrestrial poles, in order to measure optimally the galactic and solar cosmic ray flux. Energetic magnetospheric particle precipitation is monitored at lower geomagnetic latitudes. The spacecraft uses several innovative approaches including an optical fiber bus, powerful onboard computers, and large solid state memories (instead of tape recorders). Spacecraft communication and data acquisition are discussed, and the space- and ground-segment data flows are summarized. A mission lifetime of 3 years is sought with the goal of extending data acquisition over an even longer portion of the 11-year solar activity cycle.

Baker, Daniel N.↗
NASA NTRS1997

An Approach to Building a Traceability Tool for Software Development

It is difficult in a large, complex computer program to ensure that it meets the specified requirements. As the program evolves over time, a11 program constraints originally elicited during the requirements phase must be maintained. In addition, during the life cycle of the program, requirements typically change and the program must consistently reflect those changes. Imagine the following scenario. Company X wants to develop a system to automate its assembly line. With such a large system, there are many different stakeholders, e.g., managers, experts such as industrial and mechanical engineers, and end-users. Requirements would be elicited from all of the stake holders involved in the system with each stakeholder contributing their point of view to the requirements. For example, some of the requirements provided by an industrial engineer may concern the movement of parts through the assembly line. A point of view provided by the electrical engineer may be reflected in constraints concerning maximum power usage. End-users may be concerned with comfort and safety issues, whereas managers are concerned with the efficiency of the operation. With so many points of view affecting the requirements, it is difficult to manage them, communicate information to relevant stakeholders. and it is likely that conflicts in the requirements will arise. In the coding process, the implementors will make additional assumptions and interpretations on the design and the requirements of the system. During any stage of development, stakeholders may request that a requirement be added or changed. In such a dynamic environment, it is difficult to guarantee that the system will preserve the current set of requirements. Tracing, the mapping between objects in the artifacts of the system being developed, addresses this issue. Artifacts encompass documents such as the system definition, interview transcripts, memoranda, the software requirements specification, user's manuals, the functional specifications, design reports, and system code. Tracing helps 1) validate system features against, the requirement specification, 2) identify error sources and, most importantly, 3) manage change. With so many people involved in the development of the system, it becomes necessary to identify the reasons behind the design requirements or the implementation decisions. This paper is concerned with an approach that maps documents to constraints that capture properties of and relationships between the objects being modeled by the program. Section 2 provides the reader with a background on traceability tools. Section 3 gives a brief description of the context monitoring system on which the approach suggested in this paper is based. Section 4 presents an overview of our approach to providing traceability. The last section presents our future direction of research.

Delgado, Nelly↗
NASA NTRS1999

Lightweight Battery Charge Regulator Used to Track Solar Array Peak Power

A battery charge regulator based on the series-connected boost regulator (SCBR) technology has been developed for high-voltage spacecraft applications. The SCBR regulates the solar array power during insolation to prevent battery overcharge or undercharge conditions. It can also be used to provide regulated battery output voltage to spacecraft loads if necessary. This technology uses industry-standard dc-dc converters and a unique interconnection to provide size, weight, efficiency, fault tolerance, and modularity benefits over existing systems. The high-voltage SCBR shown in the photograph has demonstrated power densities of over 1000 watts per kilogram (W/kg). Using four 150-W dc-dc converter modules, it can process 2500 W of power at 120 Vdc with a minimum input voltage of 90 Vdc. Efficiency of the SCBR was 94 to 98 percent over the entire operational range. Internally, the unit is made of two separate SCBR s, each with its own analog control circuitry, to demonstrate the modularity of the technology. The analog controllers regulate the output current and incorporate the output voltage limit with active current sharing between the two units. They also include voltage and current telemetry, on/off control, and baseplate temperature sensors. For peak power tracking, the SCBR was connected to a LabView-based data acquisition system for telemetry and control. A digital control algorithm for tracking the peak power point of a solar array was developed using the principle of matching the source impedance with the load impedance for maximum energy transfer. The algorithm was successfully demonstrated in a simulated spacecraft electrical system at the Boeing PhantomWorks High Voltage Test Facility in Seattle, Washington. The system consists of a 42-string, high-voltage solar array simulator, a 77-cell, 80-ampere-hour (A-hr) nickel-hydrogen battery, and a constant power-load module. The SCBR and the LabView control algorithm successfully tracked the solar array peak power point through various load transients, including sunlight discharge transients when the total load exceeded the maximum solar array output power.

Soeder, James F.↗
NASA NTRS1981

A low-power, high-throughput maximum-likelihood convolutional decoder chip for NASA's 30/20 GHz program

It is pointed out that the NASA 30/20 GHz program will place in geosynchronous orbit a technically advanced communication satellite which can process time-division multiple access (TDMA) information bursts with a data throughput in excess of 4 GBPS. To guarantee acceptable data quality during periods of signal attenuation it will be necessary to provide a significant forward error correction (FEC) capability. Convolutional decoding (utilizing the maximum-likelihood techniques) was identified as the most attractive FEC strategy. Design trade-offs regarding a maximum-likelihood convolutional decoder (MCD) in a single-chip CMOS implementation are discussed.

Mccallister, R. D.↗
NASA NTRS2008

Phase Calibration of Antenna Arrays Aimed at Spacecraft

A document describes a method of calibrating phase differences among ground antennas in an array so that the maximum-intensity direction of the far-field interference pattern of the array coincides with the direction for aiming the antennas to enable radio communication with a distant spacecraft. The method pertains to an array typically comprising between two and four 34-m (or similar size) antennas. The antennas are first calibrated pair-wise to maximize the uplink power received at a different spacecraft that is close enough for communication via a single ground antenna. In the calibration procedure, the phase of the signal transmitted by one of the antennas is ramped through a complete cycle, thereby causing the interference pattern to sweep over this closer spacecraft and guaranteeing that, at some point during the sweep, this spacecraft is illuminated at maximum intensity. The varying received uplink power is measured by a receiver in the closer spacecraft and the measurement data are transmitted to a ground station to enable determination of the optimum phase adjustment for the direction to the closer spacecraft. This adjustment is then translated to the look direction of the distant spacecraft, which could not be reached effectively using only one antenna.

Vilnrotter, Victor↗
NASA NTRS1983

Hue-saturation-intensity split-spectrum processing of Seasat radar imagery

Characteristics of terrain features in orbital radar imagery are investigated on the basis of radar physics and the human visual system. Signal scattering, the dominant factor in radar imagery, is caused by look angle and relief displacement. Power spectra characteristics are explored, noting that surface scattering dominates the low frequencies and slope significantly affects the high frequency tone variations. The maximum human visual response has been identified for achromatic scenes as 7 cycles per degree, whereas in radar imagery the low frequency half-power point occurs near 2.5 cycles/deg. Image feature size at a scale of 1:500,000 has been calculated for the maximum human response at 1.25 km. A hue-saturation-intensity transformation is presented for displaying two or three data sets on the same image and is demonstrated for Seasat SAR imagery. The transformation is shown to restore low spatial frequencies to the SAR imagery in order to identify terrain features which would otherwise not be visually discernible.

Daily, M.↗
NASA NTRS2003

Performance Evaluation of the NEXT Ion Engine

The performance test results of three NEXT ion engines are presented. These ion engines exhibited peak specific impulse and thrust efficiency ranges of 4060 4090 s and 0.68 0.69, respectively, at the full power point of the NEXT throttle table. The performance of the ion engines satisfied all project requirements. Beam flatness parameters were significantly improved over the NSTAR ion engine, which is expected to improve accelerator grid service life. The results of engine inlet pressure and temperature measurements are also presented. Maximum main plenum, cathode, and neutralizer pressures were 12,000 Pa, 3110 Pa, and 8540 Pa, respectively, at the full power point of the NEXT throttle table. Main plenum and cathode inlet pressures required about 6 hours to increase to steady-state, while the neutralizer required only about 0.5 hour. Steady-state engine operating temperature ranges throughout the power throttling range examined were 179 303 C for the discharge chamber magnet rings and 132 213 C for the ion optics mounting ring.

Soulas, George C.↗
NASA NTRS1974

Gaseous exhaust emissions from a J-58 engine at simulated supersonic flight conditions

Emissions of total oxides of nitrogen, unburned hydrocarbons, carbon monoxide, and carbon dioxide from a J-58 engine at simulated flight conditions of Mach 2.0, 2.4, and 2.8 at 19.8 km altitude are reported. For each flight condition, measurements were made for four engine power levels from maximum power without afterburning through maximum afterburning. These measurements were made 7 cm downstream of the engine primary nozzle using a single point traversing gas sample probe. Results show that emissions vary with flight speed, engine power level, and with radial position across the exhaust.

Holdeman, J. D.↗
NASA NTRS1974

Emission calibration of a J-58 afterburning turbojet engine at simulated supersonic stratospheric flight conditions

Emissions of total oxides of nitrogen, unburned hydrocarbons, and carbon monoxide from a J-58 engine at simulated flight conditions of Mach 2.0, 2.4, and 2.8 at 19.8 km altitude are reported. For each flight condition, measurements were made for four engine power levels from maximum power without afterburning through maximum afterburning. These measurements were made 7 cm downstream of the engine primary nozzle using a single point traversing gas sample probe. Results show that emissions vary with flight speed, engine power level, and with radial position across the exhaust.

Holdeman, J. D.↗
NASA NTRS1974

Emission calibration of a J-58 afterburning turbojet engine at simulated supersonic, stratospheric flight conditions

Emissions of total oxides of nitrogen, unburned hydrocarbons, and carbon monoxide from a J-58 engine at simulated flight conditions of Mach 2.0, 2.4, and 2.8 at 19.8 km altitude are reported. For each flight condition, measurements were made for four engine power levels from maximum power without afterburning through maximum afterburning. These measurements were made 7 cm downstream of the engine primary nozzle using a single point traversing gas sample probe. Results show that emissions vary with flight speed, engine power level, and with radial position across the exhaust.

Holdeman, J. D.↗
NASA NTRS2020

STIG: A Two-Speed Transmission Aboard the Mars2020 Coring Drill

The coring drill, part of the Sampling & Caching Subsystem (SCS) aboard the Mars2020 rover, demands a wide range of drill bit torque and speed capabilities during sample acquisition operations. The two driving operating points are high speed, low torque for rotary-percussive coring, as well as low speed, high torque for separating the rock core sample from its parent rock. The spindle twin-input transmission (STIG) allows these and other operating points to be reached with an actuator of substantially less peak power and maximum current draw than that of a single-speed actuator. Rather than containing gearing of its own, the transmission interfaces to an actuator with two outputs of different gear ratios, allowing the transmission to select one of the two of the outputs to be coupled to the drill bit. This paper describes the design, capabilities, and challenges associated with the transmission and the dual-output actuator.

Szwarc, Timothy↗
NASA NTRS1982

Drag measurements in tubular structure elements. Part 3: Effect of diameter and surface structure on the drag of cylindrical tubes

Measurements on five cylinders with different surfaces show that the supercritical drag coefficient tends to 0.5 for smooth cylinders with maximum critical Re number 4.16 times 10 to the -5 power and to 0.6 for point pattern surfaces with Re number reduced to 2.16 times 10 to the -5 power. For the other surfaces, with increasing roughness the critical Re number decrease while both minimum supercritical drag coefficients increase.

Schulz, G.↗
NASA NTRSDate not supplied

Weight and Balance Considerations for Electrified Aircraft Propulsion Applied to the Parallel Electric-Gas Architecture with Synergistic Utilization Scheme (PEGASUS) Concept

The focus of this research was to understand the impact of batteries and electric motors on the weight and balance of the Parallel Electric-Gas Architecture with Synergistic Utilization Scheme (PEGASUS) concept. Because electrified aircraft propulsion components can comprise a large part of an aircraft’s weight, their integration has a significant impact on the center of gravity location and the related stability characteristics of the aircraft. We developed an analysis framework that enabled estimating the weight and volume of electrified aircraft components, their contribution to the center of gravity location, and the stability and performance of the aircraft. Trades were performed on battery placement, wing attachment point, and electric motor power to determine their impact on figures of merit such as maximum takeoff weight and block fuel. We identified batteries, and to a lesser extent electric motors, as having a large impact on aircraft center of gravity and the required horizontal tail size for longitudinal static stability. Placing electric component weight as far forward as possible resulted in reductions in the horizontal tail size required to maintain static stability. Configurations output from this framework were also evaluated using a six-degree-of-freedom simulation to quantify dynamic stability characteristics. Although shifting weight forward reduced horizontal tail size, it negatively impacted dynamic stability. This study confirms that mass property modeling and dynamic simulation, which are usually limited in conceptual design, are important for electrified aircraft concepts.

electrified aircraft propulsion↗
NASA NTRS1994

Project ECHO: Electronic Communications from Halo Orbit

The design of a communications relay to provide constant access between the Earth and the far side of the Moon is presented. Placement of the relay in a halo orbit about the L2 Earth-Moon Lagrange point allows the satellite to maintain constant simultaneous communication between Earth and scientific payloads on the far side of the Moon. The requirements of NASA's Discovery-class missions adopted and modified for this design are: total project cost should not exceed $150 million excluding launch costs, launch must be provided by Delta-class vehicle, and the satellite should maintain an operational lifetime of 10 to 15 years. The spacecraft will follow a transfer trajectory to the L2 point, after launch by a Delta II 7925 vehicle in 1999. Low-level thrust is used for injection into a stationkeeping-free halo orbit once the spacecraft reaches the L2 point. The shape of this halo orbit is highly elliptical with the maximum excursion from the L2 point being 35000 km. A spun section and despun section connected through a bearing and power transfer assembly (BAPTA) compose the structure of the spacecraft. Communications equipment is placed on the despun section to provide for a stationary dual parabolic offset-feed array antenna system. The dual system is necessary to provide communications coverage during portions of maximum excursion on the halo orbit. Transmissions to the NASA Deep Space Network 34 m antenna include six channels (color video, two voice, scientific data from lunar payloads, satellite housekeeping and telemetry and uplinked commands) using the S- and X-bands. Four radioisotope thermoelectric generators (RTG's) provide a total of 1360 W to power onboard systems and any two of the four Hughes 13 cm ion thrusters at once. Output of the ion thrusters is approximately 17.8 mN each with xenon as the propellant. Presence of torques generated by solar pressure on the antenna dish require the addition of a 'skirt' extending from the spun section of the satellite for balance. Total mass of the satellite is approximately 900 kg at a cost of $130 million FY99.

Borrelli, Jason↗
NASA NTRS1995

Maximum Acceleration Recording Circuit

Coarsely digitized maximum levels recorded in blown fuses. Circuit feeds power to accelerometer and makes nonvolatile record of maximum level to which output of accelerometer rises during measurement interval. In comparison with inertia-type single-preset-trip-point mechanical maximum-acceleration-recording devices, circuit weighs less, occupies less space, and records accelerations within narrower bands of uncertainty. In comparison with prior electronic data-acquisition systems designed for same purpose, circuit simpler, less bulky, consumes less power, costs and analysis of data recorded in magnetic or electronic memory devices. Circuit used, for example, to record accelerations to which commodities subjected during transportation on trucks.

Bozeman, Richard J., Jr.↗
NASA NTRS1968

Power consumption in acoustic amplifiers under conditions of maximum stable gain

Comparison is made of the power consumed and the acoustic amplification realized when a dc bias field is placed across a piezoelectric semiconductor and adjusted to amplify a microwave acoustic signal to the point where the forward gain is just equal to the reverse attenuation. This represents the maximum possible gain condition.

Johnson, V. R.↗
NASA NTRS2003

A Reduced Model for Prediction of Thermal and Rotational Effects on Turbine Tip Clearance

This paper describes a dynamic model that was developed to predict changes in turbine tip clearance the radial distance between the end of a turbine blade and the abradable tip seal. The clearance is estimated by using a first principles approach to model the thermal and mechanical effects of engine operating conditions on the turbine sub-components. These effects are summed to determine the resulting clearance. The model is demonstrated via a ground idle to maximum power transient and a lapse-rate takeoff transient. Results show the model demonstrates the expected pinch point behavior. The paper concludes by identifying knowledge gaps and suggesting additional research to improve the model.

Kypuros, Javier A.↗
NASA NTRS1972

Reynolds number effect on overall performance of a 10.8-centimeter (4.25-inch) sweptback bladed centrifugal compressor

The performance data were taken at 50,000 rpm, using argon gas. As the Reynolds number was reduced from near design value to 30 percent of design, the maximum efficiency decreased about 1.5 percentage points. Reducing the Reynolds number from 30 percent to approximately 10 percent of design caused the maximum efficiency to decrease another 2.5 percentage points. The variation in loss with Reynolds number is compared with inverse power relation of loss with Reynolds number.

Weigel, C.↗