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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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NASA NTRS1993

Weavability of dry polymer powder towpreg

Carbon fiber yarns (3k, 6k, 12k) were impregnated with LARC (tm) thermoplastic polyimide dry powder. Parameters for weaving these yarns were established. Eight-harness satin fabrics were successfully woven from each of the three classes of yarns and consolidated into test specimens to determine mechanical properties. It was observed that for optimum results warp yarns should have flexural rigidities between 10,000 and 100,000 mg-cm. Tow handling minimization, low tensioning, and tow bundle twisting were used to reduce fiber breakage, the separation of filaments, and tow-to-tow abrasion. No apparent effect of tow size or twist was observed on either tension or compression modulus. However, fiber damage and processing costs favor the use of 12k yarn bundles versus 3k or 6k yarn bundles in the weaving of powder-coated towpreg.

Hugh, Maylene K.↗
NASA NTRS2001

Vortical Flow Prediction Using an Adaptive Unstructured Grid Method

A computational fluid dynamics (CFD) method has been employed to compute vortical flows around slender wing/body configurations. The emphasis of the paper is on the effectiveness of an adaptive grid procedure in "capturing" concentrated vortices generated at sharp edges or flow separation lines of lifting surfaces flying at high angles of attack. The method is based on a tetrahedral unstructured grid technology developed at the NASA Langley Research Center. Two steady-state, subsonic, inviscid and Navier-Stokes flow test cases are presented to demonstrate the applicability of the method for solving practical vortical flow problems. The first test case concerns vortex flow over a simple 65' delta wing with different values of leading-edge bluntness, and the second case is that of a more complex fighter configuration. The superiority of the adapted solutions in capturing the vortex flow structure over the conventional unadapted results is demonstrated by comparisons with the windtunnel experimental data. The study shows that numerical prediction of vortical flows is highly sensitive to the local grid resolution and that the implementation of grid adaptation is essential when applying CFD methods to such complicated flow problems.

Pirzadeh, Shahyar Z.↗
NASA NTRS2001

Vortical Flow Prediction Using an Adaptive Unstructured Grid Method

A computational fluid dynamics (CFD) method has been employed to compute vortical flows around slender wing/body configurations. The emphasis of the paper is on the effectiveness of an adaptive grid procedure in "capturing" concentrated vortices generated at sharp edges or flow separation lines of lifting surfaces flying at high angles of attack. The method is based on a tetrahedral unstructured grid technology developed at the NASA Langley Research Center. Two steady-state, subsonic, inviscid and Navier-Stokes flow test cases are presented to demonstrate the applicability of the method for solving practical vortical flow problems. The first test case concerns vortex flow over a simple 65deg delta wing with different values of leading-edge bluntness, and the second case is that of a more complex fighter configuration. The superiority of the adapted solutions in capturing the vortex flow structure over the conventional unadapted results is demonstrated by comparisons with the windtunnel experimental data. The study shows that numerical prediction of vortical flows is highly sensitive to the local grid resolution and that the implementation of grid adaptation is essential when applying CFD methods to such complicated flow problems.

Pirzadeh, Shahyar Z.↗
NASA NTRS2003

Vortical Flow Prediction Using an Adaptive Unstructured Grid Method

A computational fluid dynamics (CFD) method has been employed to compute vortical flows around slender wing/body configurations. The emphasis of the paper is on the effectiveness of an adaptive grid procedure in "capturing" concentrated vortices generated at sharp edges or flow separation lines of lifting surfaces flying at high angles of attack. The method is based on a tetrahedral unstructured grid technology developed at the NASA Langley Research Center. Two steady-state, subsonic, inviscid and Navier-Stokes flow test cases are presented to demonstrate the applicability of the method for solving practical vortical flow problems. The first test case concerns vortex flow over a simple 65 delta wing with different values of leading-edge radius. Although the geometry is quite simple, it poses a challenging problem for computing vortices originating from blunt leading edges. The second case is that of a more complex fighter configuration. The superiority of the adapted solutions in capturing the vortex flow structure over the conventional unadapted results is demonstrated by comparisons with the wind-tunnel experimental data. The study shows that numerical prediction of vortical flows is highly sensitive to the local grid resolution and that the implementation of grid adaptation is essential when applying CFD methods to such complicated flow problems.

Pirzadeh, Shahyar Z.↗
NASA NTRS2005

Acoustic Predictions of Manned and Unmanned Rotorcraft Using the Comprehensive Analytical Rotorcraft Model for Acoustics (CARMA) Code System

The Comprehensive Analytical Rotorcraft Model for Acoustics (CARMA) is being developed under the Quiet Aircraft Technology Project within the NASA Vehicle Systems Program. The purpose of CARMA is to provide analysis tools for the design and evaluation of efficient low-noise rotorcraft, as well as support the development of safe, low-noise flight operations. The baseline prediction system of CARMA is presented and current capabilities are illustrated for a model rotor in a wind tunnel, a rotorcraft in flight and for a notional coaxial rotor configuration; however, a complete validation of the CARMA system capabilities with respect to a variety of measured databases is beyond the scope of this work. For the model rotor illustration, predicted rotor airloads and acoustics for a BO-105 model rotor are compared to test data from HART-II. For the flight illustration, acoustic data from an MD-520N helicopter flight test, which was conducted at Eglin Air Force Base in September 2003, are compared with CARMA full vehicle flight predictions. Predicted acoustic metrics at three microphone locations are compared for limited level flight and descent conditions. Initial acoustic predictions using CARMA for a notional coaxial rotor system are made. The effect of increasing the vertical separation between the rotors on the predicted airloads and acoustic results are shown for both aerodynamically non-interacting and aerodynamically interacting rotors. The sensitivity of including the aerodynamic interaction effects of each rotor on the other, especially when the rotors are in close proximity to one another is initially examined. The predicted coaxial rotor noise is compared to that of a conventional single rotor system of equal thrust, where both are of reasonable size for an unmanned aerial vehicle (UAV).

Boyd, D. Douglas, Jr.↗
NASA NTRS2006

High Resolution Millimeter Wave Detection of Vertical Cracks in the Space Shuttle External Tank Spray-On-Foam Insulation (SOFI)

Space Shuttle Columbia s catastrophic failure, the separation of a piece of spray-on-foam insulation (SOFI) from the external tank (ET) in the Space Shuttle Discovery s flight in 2005 and crack detected in its ET foam prior to its successful launch in 2006 emphasize the need for effective nondestructive methods for inspecting the shuttle ET SOFI. Millimeter wave nondestructive testing methods have been considered as potential and effective inspection tools for evaluating the integrity of the SOFI. This paper presents recent results of an investigation for the purpose of detecting vertical cracks in SOFI panels using a focused millimeter wave (150 GHz) reflectometer. The presented images of the SOFI panels show the capability of this reflectometer for detecting tight vertical cracks (also as a function of crack opening dimension) in exposed SOFI panels and while covered by a piece of SOFI ramp simulating a more realistic and challenging situation.

Kharkovsky, S.↗
NASA NTRS2009

Vortical Flow Prediction using an Adaptive Unstructured Grid Method

A computational fluid dynamics (CFD) method has been employed to compute vortical flows around slender wing/body configurations. The emphasis of the paper is on the effectiveness of an adaptive grid procedure in "capturing" concentrated vortices generated at sharp edges or flow separation lines of lifting surfaces flying at high angles of attack. The method is based on a tetrahedral unstructured grid technology developed at the NASA Langley Research Center. Two steady-state, subsonic, inviscid and Navier-Stokes flow test cases are presented to demonstrate the applicability of the method for solving vortical flow problems. The first test case concerns vortex flow over a simple 65 delta wing with different values of leading-edge radius. Although the geometry is quite simple, it poses a challenging problem for computing vortices originating from blunt leading edges. The second case is that of a more complex fighter configuration. The superiority of the adapted solutions in capturing the vortex flow structure over the conventional unadapted results is demonstrated by comparisons with the wind-tunnel experimental data. The study shows that numerical prediction of vortical flows is highly sensitive to the local grid resolution and that the implementation of grid adaptation is essential when applying CFD methods to such complicated flow problems.

Pirzadeh, Shahyar Z.↗
NASA NTRS2000

Analysis and Test of a Proton Exchange Membrane Fuel Cell Power System for Space Power Applications

An effort is underway to develop a prototype Proton Exchange Membrane (PEM) Fuel Cell breadboard system for fuhlre space applications. This prototype will be used to develop a comprehensive design basis for a space-rated PEM fuel cell powerplant. The prototype system includes reactant pressure regulators, ejector-based reactant pumps, a 4-kW fuel cell stack and cooling system, and a passive, membranebased oxygen / water separator. A computer model is being developed concurrently to analytically predict fluid flow in the oxidant reactant system. Fuel cells have historically played an important role in human-rated spacecraft. The Gemini and Apollo spacecraft used fuel cells for vehicle electrical power. The Space Shuttle currently uses three Alkaline Fuel Cell Powerplants (AFCP) to generate all of the vehicle's 15-20kW electrical power. Engineers at the Johnson Space Center have leveraged off the development effort ongoing in the commercial arena to develop PEM fuel cel ls for terrestrial uses. The prototype design originated from efforts to develop a PEM fuel cell replacement for the current Space Shuttle AFCP' s. In order to improve on the life and an already excellent hi storical record of reliability and safety, three subsystems were focused on. These were the fuel cell stack itself, the reactant circulation devices, and reactant / product water separator. PEM fuel cell stack performance is already demonstrating the potential for greater than four times the useful life of the current Shuttle's AFCP. Reactant pumping for product water removal has historically been accomplished with mechanical pumps. Ejectors offer an effective means of reactant pumping as well as the potential for weight reduction, control simplification, and long life. Centrifugal water separation is used on the current AFCP. A passive, membrane-based water separator offers compatibility with the micro-gravity environment of space, and the potential for control simplification, elimination of moving parts in an oxygen environment, and long life. The prototype system has been assembled from components that have previously been tested and evaluated at the component level. Preliminary data obtained from tests performed with the prototype system, as well as other published data, has been used to validate the analytical component models. These components have been incorporated into an integrated oxidant fluid system model. Results obtained from both the performance tests and the analytical model are presented.

Vasquez, Arturo↗
NASA NTRS1978

Investigation of jet exhaust and local Mach number effects on the Solid Rocket Booster (SRB) base pressures (fa19)

Two single nozzles with flare angles of 10 and 20 degrees were tested at Mach numbers of 0.5, 0.9, 1.2, 1.46, 1.96 and 3.48 in the presence of gaseous plumes. An attempt was made to determine the local Mach number above the flare by utilizing a pitot probe. This objective was only partially satisfied because the 20 degree flare separated the flow ahead of the flare for Mach numbers of 0.5 to 1.96. An accurate local Mach number could not be determined because of the separated flow. To meet the objective of a data base as a function of freestream Mach number, model surface and base pressures were obtained in the presence of gaseous plumes for a matrix of chamber pressures and temperatures at Mach numbers of 0.5, 0.9, 1.2, 1.46, 1.96 and 3.48.

Love, D. A.↗
NASA NTRS1974

Preliminary results of flight tests of vortex attenuating splines

Flight tests have been conducted to evaluate the effectiveness of a wingtip vortex attenuating device, referred to as a spline. Vortex penetrations were made with a PA-28 behind a C-54 aircraft with and without wingtip splines attached and the resultant rolling acceleration was measured and related to the roll acceleration capability of the PA-28. Tests were conducted over a range of separation distances from about 5 nautical miles (n. mi.) to less than 1 n. mi. Preliminary results indicate that, with the splines installed, there was a significant reduction in the vortex induced roll acceleration experienced by the PA-28 probe aircraft, and the distance at which the PA-28 roll control became ineffective was reduced from 2.5 n. mi. to 0.6 n. mi., or less. There was a slight increase in approach noise (approximately 4 db) with the splines installed due primarily to the higher engine power used during approach. Although splines significantly reduced the C-54 rate of climb, the rates available with four engines were acceptable for this test program. Splines did not introduce any noticeable change in the handling qualities of the C-54.

Hastings, E. C., Jr.↗
NASA NTRS1980

The effect of throat contouring on two-dimensional converging-diverging nozzles at static conditions

An experiment was conducted at static conditions to determine the internal performance effects of nozzle throat contouring, the result of increasing the circular-arc throat radius. Five nonaxisymmetric converging-diverging nozzles were tested at nozzle pressure ratios up to 9.0. Data are presented as internal thrust ratios, discharge coefficients, and static-pressure distributions. Comparisons of internal performance data for the five nozzles show that throat contouring increases the value of discharge coefficient but has no significant effect on internal thrust ratio except in cases of internal flow separation. To illustrate the use of the two dimensional converging-diverging (2-D C-D) nozzle data base, a two dimensional inviscid theory was applied to the five configurations. The generally good agreement of data with theoretical results indicates that two-dimensional inviscid theory can be successfully applied to the prediction of 2-D C-D nozzle internal flow.

Mason, M. L.↗
NASA NTRS1985

Effect of leading-edge load constraints on the design and performance of supersonic wings

A theoretical and experimental investigation was conducted to assess the effect of leading-edge load constraints on supersonic wing design and performance. In the effort to delay flow separation and the formation of leading-edge vortices, two constrained, linear-theory optimization approaches were used to limit the loadings on the leading edge of a variable-sweep planform design. Experimental force and moment tests were made on two constrained camber wings, a flat uncambered wing, and an optimum design with no constraints. Results indicate that vortex strength and separation regions were mildest on the severely and moderately constrained wings.

Darden, C. M.↗
NASA NTRS1990

Suspension systems for ground testing large space structures

A research program is documented for the development of improved suspension techniques for ground vibration testing of large, flexible space structures. The suspension system must support the weight of the structure and simultaneously allow simulation of the unconstrained rigid-body movement as in the space environment. Exploratory analytical and experimental studies were conducted for suspension systems designed to provide minimum vertical, horizontal, and rotational degrees of freedom. The effects of active feedback control added to the passive system were also investigated. An experimental suspension apparatus was designed, fabricated, and tested. This test apparatus included a zero spring rate mechanism (ZSRM) designed to support a range of weights from 50 to 300 lbs and provide vertical suspension mode frequencies less than 0.1 Hz. The lateral suspension consisted of a pendulum suspended from a moving cart (linear bearing) which served to increase the effective length of the pendulum. The torsion suspension concept involved dual pendulum cables attached from above to a pivoting support (bicycle wheel). A simple test structure having variable weight and stiffness characteristics was used to simulate the vibration characteristics of a large space structure. The suspension hardware for the individual degrees of freedom was analyzed and tested separately and then combined to achieve a 3 degree of freedom suspension system. Results from the exploratory studies should provide useful guidelines for the development of future suspension systems for ground vibration testing of large space structures.

Gold, Ronald R.↗
NASA NTRS1998

What Determines the Response: Test or Reference?

The stability of sensory memory has been studied by presenting a reference stimulus, a delay, and a test stimulus. As has been pointed out by Lages and Treisman (1998 Vision Research 38 557-572), the usual measure of performance depends only on the effect of test variations on the responses. The Weber fraction characterizing performance is more properly called the test stimulus Weber fraction. We measure the relative contribution of the test and reference to the response by the ratio of the test Weber fraction to the reference Weber fraction. The stimuli were two dark lines on a bright background. Seven reference separations, varying from 9.5 to 16.7 arc min, were intermixed in each run. Interstimulus intervals (ISI) of 50, 200 and 2000 msec and intertrial intervals (ITI) of 500 and 2500 msec were investigated. When the ISI was short (50 or 200 msec), for both ITIs, responses were determined equally by the test and reference. For the long ISI (2000 msec), the reference stimulus contributed less. However, only for the 500 msec ITI (and not for all observers) was the contribution of the reference stimulus negligible, as Treisman's criterion setting theory might suggest.

Chukova, S. V.↗
NASA NTRS2012

Initial Evaluation of a Conflict Detection Tool in the Terminal Area

Despite the recent economic recession and its adverse impact on air travel, the Federal Aviation Administration (FAA) continues to forecast an increase in air traffic demand that may see traffic double or triple by the year 2025. Increases in air traffic will burden the air traffic management system, and higher levels of safety and efficiency will be required. The air traffic controllers primary task is to ensure separation between aircraft in their airspace and keep the skies safe. As air traffic is forecasted to increase in volume and complexity [1], there is an increased likelihood of conflicts between aircraft, which adds risk and inefficiency to air traffic management and increases controller workload. To attenuate these factors, recent ATM research has shown that air and ground-based automation tools could reduce controller workload, especially if the automation is focused on conflict detection and resolution. Conflict Alert is a short time horizon conflict detection tool deployed in the Terminal Radar Approach Control (TRACON), which has limited utility due to the high number of false alerts generated and its use of dead reckoning to predict loss of separation between aircraft. Terminal Tactical Separation Assurance Flight Environment (T-TSAFE) is a short time horizon conflict detection tool that uses both flight intent and dead reckoning to detect conflicts. Results of a fast time simulation experiment indicated that TTSAFE provided a more effective alert lead-time and generated less false alerts than Conflict Alert [2]. TSAFE was previously tested in a Human-In-The-Loop (HITL) simulation study that focused on the en route phase of flight [3]. The current study tested the T-TSAFE tool in an HITL simulation study, focusing on the terminal environment with current day operations. The study identified procedures, roles, responsibilities, information requirements and usability, with the help of TRACON controllers who participated in the experiment. Metrics such as lead alert time, alert response time, workload, situation awareness and other measures were statistically analyzed. These metrics were examined from an overall perspective and comparisons between conditions (altitude resolutions via keyboard entry vs. ADS-B entry) and controller positions (two final approach sectors and two feeder sectors) were also examined. Results of these analyses and controller feedback provided evidence of T-TSAFE s potential promise as a useful air traffic controller tool. Heuristic analysis also provided information on ways in which the T-TSAFE tool can be improved. Details of analyses results will be presented in the full paper.

Verma Savita Arora↗
NASA NTRS2019

Uncertainty Analysis of the CE-22 Advanced Nozzle Test Facility

This paper presents methods and results of a detailed measurement uncertainty analysis that was performed for the Advanced Nozzle Test Facility, CE-22, located at the NASA Glenn Research Center. Results for the uncertainty in thrust and flow coefficients, in addition to other variables of interest, are provided. Results are presented separately as random uncertainty (characterizing errors inherent to instrument or measurement environment, impacting repeatability), systematic uncertainty (capturing inaccuracies due to measurement process, calibration, installation effects or other sources which may introduce bias), and total combined uncertainty. The statistical methods and engineering judgments used to estimate elemental uncertainties are described. The analysis tool MANTUS (Measurement Analysis Tool for Uncertainty in Systems) was used to quantify instrumentation system uncertainty, and statistical analysis and engineering judgment were used to quantify other random and systematic uncertainty sources. The Monte Carlo method was used to propagate systematic and random elemental uncertainties to determine the uncertainties of various calculated variables of interest.

data quality↗
NASA NTRS2017

A Sweeping Jet Application on a High Reynolds Number Semispan Supercritical Wing Configuration

The FAST-MAC circulation control model was modified to test an array of unsteady sweeping-jet actuators at realistic flight Reynolds numbers in the National Transonic Facility at the NASA Langley Research Center. Two types of sweeping jet actuators were fabricated using rapid prototype techniques, and directed over a 15% chord simple-hinged flap. The model was configured for low-speed high-lift testing with flap deflections of 30° and 60°, and a transonic cruise configuration with a 0° flap deflection. For the 30° flap high-lift configuration, the sweeping jets achieved comparable lift performance in the separation control regime, while reducing the mass flow by 54% as compared to steady blowing. However, the sweeping jets were not effective for the 60° flap. For the transonic cruise configuration, the sweeping jets reduced the drag by 3.3% at an off design condition. The drag reduction for the design lift coefficient for the sweeping jets provided only half the drag reduction shown for the steady blowing case (6.5%), but accomplished this with a 74% reduction in mass flow.

Gregory S Jones↗
NASA NTRS2019

Fusion of Surface Ceilometer Data and Satellite Cloud Retrievals in 2D Mesh Interpolating Model with Clustering

For accurate cloud ceiling information, a data fusion approach is proposed that utilizes satellite data to extend surface station information to much wider areas. Cloud base height (CBH) retrieved from satellite observations provides for much larger spatial coverage and higher resolution. The direct comparison of GOES-16 CBH with surface station ceiling yields a local bias that has to be corrected for in the initial GOES-16 cloud base information. This sparsely sampled bias correction presents an irregular 2D mesh of control points, which is then interpolated by constructing a continuous smooth field using polyharmonic splines. The influence of remote stations is restricted by grouping the control points into clusters depending on an effective distance. This cluster-based approach allows for constructing separate spline surfaces corresponding to physically different clouds. The obtained continuous bias correction function is then applied to the entire GOES-16 pixel level CBH except for areas far away from surface stations in data sparse regions such as offshore. The described method is currently being tested using daytime-only observations over the central and eastern United States. Overall, this approach has potential to provide more accurate, high spatial resolution cloud ceiling information for the aviation community.

Khlopenkov, Konstantin↗