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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 685 records · Page 38

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NASA NTRS1937

Pressure-Distribution Measurements on O-2H Airplane in Flight

Results are given of pressure-distribution measurements made over two different horizontal tail surfaces and the right wing cellule, including the slipstream area, of an observation-type biplane. Measurements were also taken of air speed, control-surface positions, control-stick forces, angular velocities, and accelerations during various abrupt maneuvers. These maneuvers consisted of push-downs and pull-ups from level flight, dive pull-outs, and aileron rolls with various thrust conditions. The results from the pressure-distribution measurements over the wing cellule are given on charts showing the variation of individual rib coefficients with wing coefficients; the data from the tail-surface pressure-distribution measurements are given mainly as total loads and moments. These data are supplemented by time histories of the measured quantities and isometric views of the rib pressure distributions occurring in abrupt maneuvers.

Pearson, H A↗

NASA NTRS1943

Wind Tunnel Tests of Ailerons at Various Speeds I : Ailerons of 0.20 Airfoil Chord and True Contour with 0.35 Aileron-chord Extreme Blunt Nose Balance on the NACA 66,2-216 Airfoil

Hinge-moment, lift, and pressure-distribution measurements were made in the two-dimensional test section of the NACA stability tunnel on a blunt-nose balance-type aileron on an NACA 66,2-216 airfoil at speeds up to 360 miles per hour corresponding to a Mach number of 0.475. The tests were made primarily to determine the effect of speed on the action of this type of aileron. The balance-nose radii of the aileron were varied from 0 to 0.02 of the airfoil chord and the gap width was varied from 0.0005 to 0.0107 of the airfoil chord. Tests were also made with the gap sealed.

FLYING, HIGH-SPEED↗

NASA NTRS1994

Martian environmental simulation for a deployable lattice mast

The Mars Pathfinder mission (formerly Mars Environmental Survey or MESUR) is scheduled for launch in December 1996 and is designed to place a small lander on the surface of Mars. After impact, the lander unfolds to expose its solar panels and release a miniature rover. Also on board is the Imager for Mars Pathfinder (IMP) binocular camera which is elevated by a deployable mast to obtain a panoramic view of the landing area. The design of this deployable mast is based on similar designs which have a long and successful flight history. In the past when this type of self-deployable mast has been used, a rate limiter has been incorporated to control the speed of deployment. In this application, to reduce weight and complexity, it was proposed to eliminate the rate limiter so that the mast would deploy without restraint. Preliminary tests showed that this type of deployment was possible especially if the deployed length was relatively short, as in this application. Compounding the problem, however, was the requirement to deploy the mast at an angle of up to 30 degrees from vertical. The deployment process was difficult to completely analyze due to the effects of gravitational and inertial loads on the mast and camera during rapid extension. Testing in a realistic manner was imperative to verify the system performance. A deployment test was therefore performed to determine the maximum tilt angle at which the mast could reliably extend and support the camera on Mars. The testing of the deployable mast requires partial gravity compensation to simulate the smaller force of Martian gravity. During the test, mass properties were maintained while weight properties were reduced. This paper describes the testing of a deployable mast in a simulated Martian environment as well as the results of the tests.

Warden, Robert M.↗

NASA NTRS1995

Comparing Ulysses wind speed with coronal flux-tube expansion factor

We have been comparing measurements of solar wind speed at the Ulysses spacecraft with coronal flux-tube expansion rates, derived from photospheric field measurements using a current-free coronal model. The large-scale patterns of derived speed have continued to reproduce the observed patterns from launch through south polar passage to the present 40S latitude of the spacecraft. The fastest non-transient wind speeds of approx. 860 km/s were encountered at midlatitudes en route to the south pole, rather than during polar passage when the peak speeds were approx. 820 km/s. Although this result is in qualitative agreement with the idea that the wind speed is controlled by the coronal flux-tube expansion rate, the 40 km/s difference is significantly smaller than the 100-150 km/s difference based on our in-ecliptic calibration. This paper will summarize our attempts to resolve this discrepancy and will show the observational status of our coronal/interplanetary comparison at the time of the meeting.

Sheeley, N. R., Jr.↗

NASA NTRS1995

Spectral Changes in Metal Halide and High-pressure Sodium Lamps Equipped with Electronic Dimming

Electronic dimming of high-intensity discharge lamps offers control of photosynthetic photon flux (PPF) but is often characterized as causing significant spectral changes. Growth chambers with 400-W metal halide (MH) and high-pressure sodium (HPS) lamps were equipped with a dimmer system using silicon-controlled rectifiers (SCR) as high-speed switches. Phase control operation turned the line power off for some period of the alternating current cycle. At full power, the electrical input to HPS and MH lamps was 480 W (root mean squared) and could be decreased to 267 W and 428 W, respectively, before the arc was extinguished. Concomitant with this decrease in input power, PPF decreased by 60% in HPS and 50% in MH. The HPS lamp has characteristic spectral peaks at 589 and 595 nm. As power to the HPS lamps was decreased, the 589-nm peak remained constant while the 595-nm peak decreased, equaling the 589-nm peak at 345-W input, and the 589-nm peak was almost absent at 270-W input. The MH lamp has a broader spectral output but also has a peak at 589 nm and another smaller peak at 545 nm. As input power to the MH lamps decreased, the peak at 589 diminished to equal the 545-nm peak. As input power approached 428 W, the 589-nm peak shifted to 570 nm. While the spectrum changed as input power was decreased in the MH and HPS lamps, the phytochrome equilibrium ratio (P(sub fr):P(sub tot)) remains unchanged for both lamp types.

Bubenheim, David L.↗

NASA NTRS1995

Spectral Changes in Metal Halide and High-Pressure Sodium Lamps Equipped with Electronic Dimming

Electronic dimming of high-intensity discharge lamps offers control of Photosynthetic Photon Flux (PPF) but is often characterized as causing significant spectral changes. Growth chambers with 400-W Metal Halide (MH) and High-Pressure Sodium (HPS) lamps were equipped with a dimmer system using Silicon-Controlled Rectifiers (SCR) as high-speed switches. Phase control operation turned the line power off for some period of the alternating current cycle. At full power, the electrical input to HPS and MH lamps was 480 W (root mean squared) and could be decreased to 267 W and 428 W, respectively, before the arc was extinguished. Concomitant with this decrease in input power, PPF decreased by 60% in HPS and 50% in MH. The HPS lamp has characteristic spectral peaks at 589 and 595 nm. As power to the HPS lamps was decreased, the 589-nm peak remained constant while the 595-nm peak decreased, equaling the 589-nm peak at 345-W input, and 589-nm peak was almost absent at 270-W input. The MH lamp has a broader spectral output but also has a peak at 589 nm and another smaller peak at 545 nm. As input power approached 428 W, the 589-nm peak shifted to 570 nm. While the spectrum changed as input power was decreased in the MH and HPS lamps, the phytochrome equilibrium ratio (P(sub ft):P(sub tot)) remains unchanged for both lamp types.

Bubenheim, David L.↗

NASA NTRS2000

Solar Powered Aircraft, Photovoltaic Array/Battery System Tabletop Demonstration: Design and Operation Manual

A system was constructed to demonstrate the power system operation of a solar powered aircraft. The system consists of a photovoltaic (PV) array, a charge controller, a battery, an electric motor and propeller. The system collects energy from the PV array and either utilizes this energy to operate an electric motor or stores it in a rechargeable battery for future use. The system has a control panel which displays the output of the array and battery as well as the total current going to the electric motor. The control panel also has a means for adjusting the output to the motor to control its speed. The entire system is regulated around 12 VDC.

Colozza, Anthony J.↗

NASA NTRS1994

Initial Field Evaluation of Pilot Procedures for Flying CTAS Descent Clearances

The Center TRACON Automation System (CTAS) is a new support system that is designed to assist air traffic controllers in the management of arrival traffic. CTAS will provide controllers with more information about current air traffic, enabling them to provide clearances for efficient, conflict-free descents that help achieve an orderly stream of aircraft at the final approach fix. CTAS is a computer-based system that functions as a "ground-based FMS" that can predict flight trajectories and arrival times for all incoming aircraft. CTAS uses an aircraft's cruise airspeed; current air traffic, winds and temperature; performance characteristics of the aircraft type; and individual airline preferences to create a flight profile from cruise altitude to the final approach fix. Controllers can use this flight profile to provide a descent clearance that will allow an aircraft to fly an efficient descent and merge more smoothly with other arriving aircraft. A field test of the CTAS Descent Advisor software was conducted at the Denver Center for aircraft arriving at the Stapleton International Airport from September 12-29. CTAS Descent clearances were given to a NASA flight test aircraft and to 77 airline flights that arrived during low traffic periods. For the airline portion of the field test, cockpit procedures and pilot briefing packages for both FMS equipped and unequipped aircraft were developed in cooperation with an airline. The procedures developed for the FMS equipped aircraft were to fly a VNAV descent at a controller specified speed to cross a metering fix at a specified altitude and speed. For nonFMS aircraft, the clearance also specified a CTAS calculated top-of-descent point. Some CTAS related flight deck issues included how much time was available to the pilots' for compliance, the amount of information that needed to be interpreted in the clearance and possible repercussions of misunderstandings. Data collected during the study ranged from subjective data (including the airline pilots' opinions and comments about the new descent clearances and procedures) to objective data (including observations of aircraft performance from the flight deck). This paper will present data and the resulting changes in the design of the procedures and clearance phraseology.

Palmer, Everett↗

NASA NTRS1997

Unique Tuft Test Facility Dramatically Reduces Brush Seal Development Costs

Brush seals have been incorporated in the latest turbine engines to reduce leakage and improve efficiency. However, the life of these seals is limited by wear. Studies have shown that optimal sealing characteristics for a brush seal occur before the interference fit between the brush and shaft is excessively worn. Research to develop improved tribopairs (brush and coating) with reduced wear and lower friction has been hindered by the lack of an accurate, low-cost, efficient test methodology. Estimated costs for evaluating a new material combination in an engine company seal test program are on the order of $100,000. To address this need, the NASA Lewis Research Center designed, built, and validated a unique, innovative brush seal tuft tester that slides a single tuft of brush seal wire against a rotating shaft under controlled loads, speeds, and temperatures comparable to those in turbine engines. As an initial screening tool, the brush seal tuft tester can tribologicaly evaluate candidate seal materials for 1/10th the cost of full-scale seal tests. Previous to the development of the brush seal tuft tester facility, most relevant tribological data had been obtained from full-scale seal tests conducted primarily to determine seal leakage characteristics. However, from a tribological point of view, these tests included the confounding effects of varying contact pressures, bristle flaring, high-temperature oxidation, and varying bristle contact angles. These confounding effects are overcome in tuft testing. The interface contact pressures can be either constant or varying depending on the tuft mounting device, and bristle wear can be measured optically with inscribed witness marks. In a recent cooperative program with a U.S. turbine engine manufacturer, five metallic wire candidates were tested against a plasma-sprayed Nichrome-bonded chrome carbide. The wire materials used during this collaboration were either nickel-chrome- or cobaltchrome-based superalloys. These tests corroborated full-scale seal test results and provided insight into previously untested combinations. As the cycle temperature for improved efficiency turbine engines increases, new brush seal materials combinations must be considered. Future brush seal tuft testing will include both metallic and ceramic bristles versus commercial and NASA-developed shaft coatings. The ultimate goal of this work is to expand the current data base so that seal designers can tailor brush seal materials to specific applications.

Fellenstein, James A.↗

NASA NTRS1998

Parallel processing spacecraft communication system

An uplink controlling assembly speeds data processing using a special parallel codeblock technique. A correct start sequence initiates processing of a frame. Two possible start sequences can be used; and the one which is used determines whether data polarity is inverted or non-inverted. Processing continues until uncorrectable errors are found. The frame ends by intentionally sending a block with an uncorrectable error. Each of the codeblocks in the frame has a channel ID. Each channel ID can be separately processed in parallel. This obviates the problem of waiting for error correction processing. If that channel number is zero, however, it indicates that the frame of data represents a critical command only. That data is handled in a special way, independent of the software. Otherwise, the processed data further handled using special double buffering techniques to avoid problems from overrun. When overrun does occur, the system takes action to lose only the oldest data.

Bolotin, Gary S.↗

NASA NTRS2008

Focused Assessment of State-of-the-Art CFD Capabilities for Prediction of Subsonic Fixed Wing Aircraft Aerodynamics

Several recent workshops and studies are used to make an assessment of the current status of CFD for subsonic fixed wing aerodynamics. Uncertainty quantification plays a significant role in the assessment, so terms associated with verification and validation are given and some methodology and research areas are highlighted. For high-subsonic-speed cruise through buffet onset, the series of drag prediction workshops and NASA/Boeing buffet onset studies are described. For low-speed flow control for high lift, a circulation control workshop and a synthetic jet flow control workshop are described. Along with a few specific recommendations, gaps and needs identified through the workshops and studies are used to develop a list of broad recommendations to improve CFD capabilities and processes for this discipline in the future.

Rumsey, Christopher L.↗

NASA NTRS2007

Interface for Physics Simulation Engines

DSS-Prototyper is an open-source, realtime 3D virtual environment software that supports design simulation for the new Vision for Space Exploration (VSE). This is a simulation of NASA's proposed Robotic Lunar Exploration Program, second mission (RLEP2). It simulates the Lunar Surface Access Module (LSAM), which is designed to carry up to four astronauts to the lunar surface for durations of a week or longer. This simulation shows the virtual vehicle making approaches and landings on a variety of lunar terrains. The physics of the descent engine thrust vector, production of dust, and the dynamics of the suspension are all modeled in this set of simulations. The RLEP2 simulations are drivable (by keyboard or joystick) virtual rovers with controls for speed and motor torque, and can be articulated into higher or lower centers of gravity (depending on driving hazards) to enable drill placement. Gravity also can be set to lunar, terrestrial, or zero-g. This software has been used to support NASA's Marshall Space Flight Center in simulations of proposed vehicles for robotically exploring the lunar surface for water ice, and could be used to model all other aspects of the VSE from the Ares launch vehicles and Crew Exploration Vehicle (CEV) to the International Space Station (ISS). This simulator may be installed and operated on any Windows PC with an installed 3D graphics card.

Damer, Bruce↗

NASA NTRS2013

Integrating a Microwave Radiometer into Radar Hardware for Simultaneous Data Collection Between the Instruments

The conventional method for integrating a radiometer into radar hardware is to share the RF front end between the instruments, and to have separate IF receivers that take data at separate times. Alternatively, the radar and radiometer could share the antenna through the use of a diplexer, but have completely independent receivers. This novel method shares the radar's RF electronics and digital receiver with the radiometer, while allowing for simultaneous operation of the radar and radiometer. Radars and radiometers, while often having near-identical RF receivers, generally have substantially different IF and baseband receivers. Operation of the two instruments simultaneously is difficult, since airborne radars will pulse at a rate of hundreds of microseconds. Radiometer integration time is typically 10s or 100s of milliseconds. The bandwidth of radar may be 1 to 25 MHz, while a radiometer will have an RF bandwidth of up to a GHz. As such, the conventional method of integrating radar and radiometer hardware is to share the highfrequency RF receiver, but to have separate IF subsystems and digitizers. To avoid corruption of the radiometer data, the radar is turned off during the radiometer dwell time. This method utilizes a modern radar digital receiver to allow simultaneous operation of a radiometer and radar with a shared RF front end and digital receiver. The radiometer signal is coupled out after the first down-conversion stage. From there, the radar transmit frequencies are heavily filtered, and the bands outside the transmit filter are amplified and passed to a detector diode. This diode produces a DC output proportional to the input power. For a conventional radiometer, this level would be digitized. By taking this DC output and mixing it with a system oscillator at 10 MHz, the signal can instead be digitized by a second channel on the radar digital receiver (which typically do not accept DC inputs), and can be down-converted to a DC level again digitally. This unintuitive step allows the digital receiver to sample both the radiometer and radar data at a rapid, synchronized data rate (greater than 1 MHz bandwidth). Once both signals are sampled by the same digital receiver, high-speed quality control can be performed on the radiometer data to allow it to take data simultaneously with the radar. The radiometer data can be blanked during radar transmit, or when the radar return is of a power level high enough to corrupt the radiometer data. Additionally, the receiver protection switches in the RF front end can double as radiometer calibration sources, the short (four-microsecond level) switching periods integrated over many seconds to estimate the radiometer offset. The major benefit of this innovation is that there is minimal impact on the radar performance due to the integration of the radiometer, and the radiometer performance is similarly minimally affected by the radar. As the radar and radiometer are able to operate simultaneously, there is no extended period of integration time loss for the radiometer (maximizing sensitivity), and the radar is able to maintain its full number of pulses (increasing sensitivity and decreasing measurement uncertainty).

McLinden, Matthew↗

NASA NTRS2015

Air Traffic Management Technology Demostration: 1 Research and Procedural Testing of Routes

NASA's Air Traffic Management Technology Demonstration-1 (ATD-1) will operationally demonstrate the feasibility of efficient arrival operations combining ground-based and airborne NASA technologies. The ATD-1 integrated system consists of the Traffic Management Advisor with Terminal Metering which generates precise time-based schedules to the runway and merge points; Controller Managed Spacing decision support tools which provide controllers with speed advisories and other information needed to meet the schedule; and Flight deck-based Interval Management avionics and procedures which allow flight crews to adjust their speed to achieve precise relative spacing. Initial studies identified air-ground challenges related to the integration of these three scheduling and spacing technologies, and NASA's airborne spacing algorithm was modified to address some of these challenges. The Research and Procedural Testing of Routes human-in-the-loop experiment was then conducted to assess the performance of the new spacing algorithm. The results of this experiment indicate that the algorithm performed as designed, and the pilot participants found the airborne spacing concept, air-ground procedures, and crew interface to be acceptable. However, the researchers concluded that the data revealed issues with the frequency of speed changes and speed reversals.

Wilson, Sara R.↗

NASA NTRS2023

Strategic Deconfliction of Small Unmanned Aircraft Using Operational Volume Blocks at Crossing Waypoints

In this research, first, analytical case studies are performed to understand the parameters on which the minimum temporal separation between unmanned aircraft at crossing waypoints is dependent for enabling strategic deconfliction. The analytical expressions show that the minimum temporal separation is a function of the length and width of operational volume blocks, the relative positions of the active operational volume blocks, groundspeed of unmanned aircraft, and the incoming crossing angle. Next, the parametric study shows that the impact of the incoming crossing angle on the minimum temporal separation at a crossing waypoint increases with an increase in the width of the operational volume blocks. Finally, simulation studies are performed to understand the impact of operational volume block sizing, on-demand departure rate, minimum departure time separation, and the incoming crossing angle on the average ground delay of unmanned aircraft traveling on two routes with a single crossing waypoint and identical on-demand departure rate. Each unmanned aircraft’s estimated time of arrival at a crossing waypoint is adjusted by introducing a ground delay in departure time; no other controls (e.g., speed adjustments) are applied for strategic deconfliction. Simulation studies show that the impact of the minimum temporal separation at a crossing waypoint on the average ground delay of flights is negligible if the minimum departure time separation is at least two times the minimum temporal separation. Therefore, with an increase in minimum departure time separation at a depot, the impact of increased length of operational volume blocks enclosing the crossing waypoint on the ground delay is offset to an extent.

Operational intent↗

NASA NTRS2023

Strategic Deconfliction of Small Unmanned Aircraft Using Operational Volume Blocks at Crossing Waypoints

Operational Intent↗

NASA NTRSDate not supplied

Designing A Coaxial Quadrotor for Urban Air Mobility

The NASA Revolutionary Vertical Lift Technology Project has created several Urban Air Mobility reference design vehicles to aid the burgeoning UAM industry. With over 700 conceptual designs already proposed, much interest has been generated across many unique configurations to bring vertical flight to the masses. The quadrotor configuration is one of the NASA reference vehicles, and does have some conceptual designs proposed in the industry. This study looks at the conceptual design of a coaxial quadrotor vehicle as an extension of the NASA RVLT quadrotor concept vehicle. The coaxial quadrotor has a similar layout to the quadrotor, but with a coaxial rotor at each of the four corners of the vehicle. Trade studies carried out using the NASA Design and Analysis of Rotorcraft (NDARC) tool analyzed the vehicle sizing, design gross weight, and performance for four design variants. The variants all use the baseline coaxial quadrotor layout, but analyze the effects of two major design choices: 1) variable pitch vs variable speed rotor control, and 2) motor to rotor power transmission via a gearbox vs direct-drive. The results indicate one design variant may have more benefits on some of the design objectives while another configuration may better satisfy the others.

UAM↗

NASA NTRSDate not supplied

High Reynolds Number Testing of the NATO AVT-298 SWiFT Configuration at the National Transonic Facility

A high Reynolds number wind tunnel test of the NATO AVT-298 Swept Wing Flow Test (SWiFT) configuration was conducted in the National Transonic Facility at the NASA Langley Research Center during the summer of 2023. The SWiFT research model geometry has relevance to both blended/hybrid wing body (BWB/HWB) and unmanned combat aerial vehicle (UCAV) configurations, and the test campaign was the culmination of an international collaboration under the NATO AVT-298 research task group. The main objectives of the test were to investigate Reynolds number scaling effects on low-speed stability & control characteristics and to examine the onset and progression of flow separation on the wings particularly near the wing crank. Force & moment and surface pressure data were acquired at Mach numbers from 0.2 to 0.8, Reynolds numbers from 2.5 to 34 million, angles of attack from -3 to 20 degrees, and sideslip angles from -10 to 10 degrees. Boundary layer transition detection techniques utilizing static pressure taps, unsteady pressure transducers, and sublimating chemicals were used on the model in a free/natural transition state or a forced transition state using trip dots. Pressure sensitive paint was used to obtain a global surface pressure profile on the model and an advanced laser velocimetry technique was used to obtain velocity measurements in the wake downstream of the wing crank. The results from the test showed clear Reynolds number scaling effects on the maximum lift coefficient and pitching moment coefficient at Mach 0.2, while also capturing significant hysteresis effects. The data acquired from the test will ultimately help improve computational aerodynamic analysis and design tools for application to future BWB-type vehicle configurations.

NATO AVT-298↗