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At least 163 records · Page 9

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NASA NTRS2007

Situational Leadership in Air Traffic Control

In high-risk environments such as air traffic control, leadership on different levels plays a certain role in establishing, promoting, and maintaining a good safety culture. The current study aimed to investigate how leadership styles, leadership style adaptability, and over and under task leadership behavior differed across situations, operative conditions, leadership structures, and working tasks in an air traffic control setting. Study locations were two air traffic control centers in Sweden with different operational conditions and leadership structures, and an administrative air traffic management unit. Leadership was measured with a questionnaire based on Leader Effectiveness and Adaptability Description (LEAD; Blanchard, Zigarmi & Zigarmi, 2003; Hersey & Blanchard, 1988). The results showed that the situation had strong impact on the leadership in which the leadership behavior was more relationship oriented in Success and Group situations than in Hardship and Individual situations. The leadership adaptability was further superior in Success and Individual situations compared with Hardship and Group situations. Operational conditions, leadership structures and working tasks were, on the other hand, not associated with leadership behavior.

Arvidsson, Marcus↗

NASA NTRS1989

Real-time closed-loop simulation and upset evaluation of control systems in harsh electromagnetic environments

Digital control systems for applications such as aircraft avionics and multibody systems must maintain adequate control integrity in adverse as well as nominal operating conditions. For example, control systems for advanced aircraft, and especially those with relaxed static stability, will be critical to flight and will, therefore, have very high reliability specifications which must be met regardless of operating conditions. In addition, multibody systems such as robotic manipulators performing critical functions must have control systems capable of robust performance in any operating environment in order to complete the assigned task reliably. Severe operating conditions for electronic control systems can result from electromagnetic disturbances caused by lightning, high energy radio frequency (HERF) transmitters, and nuclear electromagnetic pulses (NEMP). For this reason, techniques must be developed to evaluate the integrity of the control system in adverse operating environments. The most difficult and illusive perturbations to computer-based control systems that can be caused by an electromagnetic environment (EME) are functional error modes that involve no component damage. These error modes are collectively known as upset, can occur simultaneously in all of the channels of a redundant control system, and are software dependent. Upset studies performed to date have not addressed the assessment of fault tolerant systems and do not involve the evaluation of a control system operating in a closed-loop with the plant. A methodology for performing a real-time simulation of the closed-loop dynamics of a fault tolerant control system with a simulated plant operating in an electromagnetically harsh environment is presented. In particular, considerations for performing upset tests on the controller are discussed. Some of these considerations are the generation and coupling of analog signals representative of electromagnetic disturbances to a control system under test, analog data acquisition, and digital data acquisition from fault tolerant systems. In addition, a case study of an upset test methodology for a fault tolerant electromagnetic aircraft engine control system is presented.

Belcastro, Celeste M.↗

NASA NTRSDate not supplied

Extended Wear Testing of the 12-kW Advanced Electric Propulsion System Engineering Test Unit Hall Thruster

This work presents a summary of the first wear test of the 12 kW Advanced Electric Propulsion System (AEPS) intended to isolate the impact of discharge current on component wear rates. Testing was led by Aerojet Rocketdyne and performed with the AEPS Engineering Test Unit 2 (ETU-2) thruster. In total, approximately 860 hours of operation were accumulated split between operating conditions of 600 V/9 kW, 600 V/11 kW, and 600 V/12 kW. Thruster performance and stability were invariant throughout the wear test for all thruster throttle conditions and shown to be equal to the values previously measured with ETU-2. Inner front pole cover erosion rates were found to be invariant to discharge current as the measurements at all three operating conditions were equal to within the empirical uncertainty. Outer front pole cover erosion rates were found to be equal for the 600 V/9 kW and 600 V/12 kW conditions with operation at 600 V/11 kW yielding elevated erosion rates. Overall, the data shows that the AEPS thruster design has a high probability of meeting lifetime and performance requirements as the design proceeds to flight production and qualification.

Jason D Frieman↗

NASA NTRSDate not supplied

Extended Wear Testing of the 12-kW Advanced Electric Propulsion System Engineering Test Unit Hall Thruster

This work presents a summary of the first wear test of the 12 kW Advanced Electric Propulsion System (AEPS) intended to isolate the impact of discharge current on component wear rates. Testing was led by Aerojet Rocketdyne and performed with the AEPS Engineering Test Unit 2 (ETU-2) thruster. In total, approximately 860 hours of operation were accumulated split between operating conditions of 600 V/9 kW, 600 V/11 kW, and 600 V/12 kW. Thruster performance and stability were invariant throughout the wear test for all thruster throttle conditions and shown to be equal to the values previously measured with ETU-2. Inner front pole cover erosion rates were found to be invariant to discharge current as the measurements at all three operating conditions were equal to within the empirical uncertainty. Outer front pole cover erosion rates were found to be equal for the 600 V/9 kW and 600 V/12 kW conditions with operation at 600 V/11 kW yielding rates that were comparable to those on the inner front pole cover. Overall, the data shows that the AEPS thruster design has a high probability of meeting lifetime and performance requirements as the design proceeds to flight production and qualification.

HERMeS↗

NASA NTRS2002

RWPV bioreactor mass transport: earth-based and in microgravity

Mass transport and mixing of perfused scalar quantities in the NASA Rotating Wall Perfused Vessel bioreactor are studied using numerical models of the flow field and scalar concentration field. Operating conditions typical of both microgravity and ground-based cell cultures are studied to determine the expected vessel performance for both flight and ground-based control experiments. Results are presented for the transport of oxygen with cell densities and consumption rates typical of colon cancer cells cultured in the RWPV. The transport and mixing characteristics are first investigated with a step change in the perfusion inlet concentration by computing the time histories of the time to exceed 10% inlet concentration. The effects of a uniform cell utilization rate are then investigated with time histories of the outlet concentration, volume average concentration, and volume fraction starved. It is found that the operating conditions used in microgravity produce results that are quite different then those for ground-based conditions. Mixing times for microgravity conditions are significantly shorter than those for ground-based operation. Increasing the differential rotation rates (microgravity) increases the mixing and transport, while increasing the mean rotation rate (ground-based) suppresses both. Increasing perfusion rates enhances mass transport for both microgravity and ground-based cases, however, for the present range of operating conditions, above 5-10 cc/min there are diminishing returns as much of the inlet fluid is transported directly to the perfusion exit. The results show that exit concentration is not a good indicator of the concentration distributions in the vessel. In microgravity conditions, the NASA RWPV bioreactor with the viscous pump has been shown to provide an environment that is well mixed. Even when operated near the theoretical minimum perfusion rates, only a small fraction of the volume provides less than the required oxygen levels. 2002 Wiley Periodicals, Inc.

Models, Biological↗

NASA NTRS2004

Probabilistic Study Conducted on Sensor-Based Engine Life Calculation

Turbine engine life management is a very complicated process to ensure the safe operation of an engine subjected to complex usage. The challenge of life management is to find a reasonable compromise between the safe operation and the maximum usage of critical parts to reduce maintenance costs. The commonly used "cycle count" approach does not take the engine operation conditions into account, and it oversimplifies the calculation of the life usage. Because of the shortcomings, many engine components are regularly pulled for maintenance before their usable life is over. And, if an engine has been running regularly under more severe conditions, components might not be taken out of service before they exceed their designed risk of failure. The NASA Glenn Research Center and its industrial and academic partners have been using measurable parameters to improve engine life estimation. This study was based on the Monte Carlo simulation of 5000 typical flights under various operating conditions. First a closed-loop engine model was developed to simulate the engine operation across the mission profile and a thermomechanical fatigue (TMF) damage model was used to calculate the actual damage during takeoff, where the maximum TMF accumulates. Next, a Weibull distribution was used to estimate the implied probability of failure for a given accumulated cycle count. Monte Carlo simulations were then employed to find the profiles of the TMF damage under different operating assumptions including parameter uncertainties. Finally, probabilities of failure for different operating conditions were analyzed to demonstrate the importance of a sensor-based damage calculation in order to better manage the risk of failure and on-wing life.

Guo, Ten-Huei↗

NASA NTRS1979

Experimental clean combustor program: Diesel no. 2 fuel addendum, phase 3

A CF6-50 engine equipped with an advanced, low emission, double annular combustor was operated 4.8 hours with No. 2 diesel fuel. Fourteen steady-state operating conditions ranging from idle to full power were investigated. Engine/combustor performance and exhaust emissions were obtained and compared to JF-5 fueled test results. With one exception, fuel effects were very small and in agreement with previously obtained combustor test rig results. At high power operating condition, the two fuels produced virtually the same peak metal temperatures and exhaust emission levels. At low power operating conditions, where only the pilot stage was fueled, smoke levels tended to be significantly higher with No. 2 diesel fuel. Additional development of this combustor concept is needed in the areas of exit temperature distribution, engine fuel control, and exhaust emission levels before it can be considered for production engine use.

Gleason, C. C.↗

NASA NTRS2001

Experimental Assessment of the Emissions Control Potential of a Rich/Quench/Lean Combustor for High Speed Civil Transport Aircraft Engines

In support of Pratt & Whitney efforts to define the Rich burn/Quick mix/Lean burn (RQL) combustor for the High Speed Civil Transport (HSCT) aircraft engine, UTRC conducted a flametube-scale study of the RQL concept. Extensive combustor testing was performed at the Supersonic Cruise (SSC) condition of a HSCT engine cycle, Data obtained from probe traverses near the exit of the mixing section confirmed that the mixing section was the critical component in controlling combustor emissions. Circular-hole configurations, which produced rapidly-, highly-penetrating jets, were most effective in limiting NOx. The spatial profiles of NOx and CO at the mixer exit were not directly interpretable using a simple flow model based on jet penetration, and a greater understanding of the flow and chemical processes in this section are required to optimize it. Neither the rich-combustor equivalence ratio nor its residence time was a direct contributor to the exit NOx. Based on this study, it was also concluded that (1) While NOx formation in both the mixing section and the lean combustor contribute to the overall emission, the NOx formation in the mixing section dominates. The gas composition exiting the rich combustor can be reasonably represented by the equilibrium composition corresponding to the rich combustor operating condition. Negligible NOx exits the rich combustor. (2) At the SSC condition, the oxidation processes occurring in the mixing section consume 99 percent of the CO exiting the rich combustor. Soot formed in the rich combustor is also highly oxidized, with combustor exit SAE Smoke Number <3. (3) Mixing section configurations which demonstrated enhanced emissions control at SSC also performed better at part-power conditions. Data from mixer exit traverses reflected the expected mixing behavior for off-design jet to crossflow momentum-flux ratios. (4) Low power operating conditions require that the RQL combustor operate as a lean-lean combustor to achieve low CO and high efficiency. (5) A RQL combustor can achieve the emissions goal of EINOX = 5 at the Supersonic Cruise operating condition for a HSCT engine.

Rosfjord, T. J.↗

NASA NTRS2001

Experimental Assessment of the Emissions Control Potential of a Rich/Quench/ Lean Combustor for High Speed Civil Transport Aircraft Engines

In support of Pratt & Whitney efforts to define the Rich burn/Quick mix/Lean burn (RQL) combustor for the High Speed Civil Transport (HSCT) aircraft engine, UTRC conducted a flametube-scale study of the RQL concept. Extensive combustor testing was performed at the Supersonic Cruise (SSC) condition of an HSCT engine cycle. Data obtained from probe traverses near the exit of the mixing section confirmed that the mixing section was the critical component in controlling combustor emissions. Circular-hole configurations, which produced rapidly-, highly-penetrating jets, were most effective in limiting NO(x). The spatial profiles of NO(x) and CO at the mixer exit were not directly interpretable using a simple flow model based on jet penetration, and a greater understanding of the flow and chemical processes in this section are required to optimize it. Neither the rich-combustor equivalence ratio nor its residence time was a direct contributor to the exit NO(x). Based on this study, it was also concluded that: (1) While NO(x) formation in both the mixing section and the lean combustor contribute to the overall emission, the NOx formation in the mixing section dominates. The gas composition exiting the rich combustor can be reasonably represented by the equilibrium composition corresponding to the rich combustor operating condition. Negligible NO(x) exits the rich combustor. (2) At the SSC condition, the oxidation processes occurring in the mixing section consume 99 percent of the CO exiting the rich combustor. Soot formed in the rich combustor is also highly oxidized, with combustor exit SAE Smoke Number <3. (3) Mixing section configurations which demonstrated enhanced emissions control at SSC also performed better at part-power conditions. Data from mixer exit traverses reflected the expected mixing behavior for off-design jet to crossflow momentum-flux ratios. (4) Low power operating conditions require that the RQL combustor operate as a lean-lean combustor to achieve low CO and high efficiency. (5) An RQL combustor can achieve the emissions goal of EINO(x) = 5 at the Supersonic Cruise operating condition for an HSCT engine.

Tacina, Robert R.↗

NASA NTRS2023

Ventilation Heat Exchanger / Flow Meter for xPLSS

The flow meter / heat exchanger (FMHX) in the ventilation loop of the exploration EMU cools the ventilation gas and measures the ventilation flow rate. The heat exchanger transfers heat from the ventilation gas to the thermal control loop via a miniature shell-and-tube heat exchanger. The flow meters calculate the flow rate of gas through the ventilation loop based on the pressure drop across the heat exchanger core. Creare has delivered four design validation test (DVT) heat exchangers and five DVT flow meters to NASA JSC to support development of the exploration portable life support system (xPLSS). This paper describes the design and performance of the DVT units.The heat exchangers are designed to cool the ventilation gas to a specified temperature with low pressure losses under the most challenging operating conditions. The measured performance of the DVT heat exchangers agrees well with design models and meets all performance requirements. The flow meters use a MEMS thermal flow sensor to produce a signal that is proportional to a small bypass flow around the heat exchanger core. They are designed to achieve high measurement accuracy across the full range of xPLSS operating conditions. We calibrated the flow meters in a special-purpose flow facility that simulates operation in the xPLSS ventilation loop. Calibration testing shows that DVT flow meters produce digital output for vent loop mass flow that meets NASA’s accuracy requirements across the range of xPLSS operating conditions. This paper reviews the design of the heat exchangers and flow meters and presents data from the final flow meter calibration testing, heat exchanger performance validation, and initial ground testing in NASA’s xPLSS.

spacesuit↗

NASA NTRS2023

Ventilation Heat Exchanger / Flow Meter for xPLSS

spacesuit↗

NASA NTRS2018

Investigations of Pole Erosion Mechanisms in the 12.5 kW HERMeS Hall Thruster Using Numerical Simulations and Ion Velocity Measurements

When the Hall Effect Rocket with Magnetic Shielding (HERMeS) was tested at the NASA GRC’s VF-5 the erosion rate on the pole covers was found to be at least a factor of two higher at 300 V and 20.8 A than at other operating conditions in the 400 V to 600 V range. Simulations using our hydrodynamics code Hall2De do not predict increased erosion rates at 300 V but accurately compute erosion rates similar to the measurements at the other operating conditions. We investigate the source of the discrepancy between measurements and simulations at 300 V using a combination of numerical simulations and experimental measurements of the ion velocity fields in the acceleration region of the thruster. By examining previous simulations that predicted the measured erosion rates for other thrusters and operating conditions, we determine that sputtering of the pole surfaces by high energy ions is the most likely mechanism behind the erosion rates at 300 V. High energy ions sputter the pole surfaces when the acceleration region of the thruster is downstream of the pole surface plane. When the latter occurs, the curvature of the plasma potential contours at the edges of the channel accelerates a small fraction of the high energy ions radially. We find that neither our Hall2De simulations nor the experimental measurements produce ions of sufficiently high energy and flux to the inner pole cover to explain the measured erosion there. Only at the outer edge of the cover is where we find the ions needed to yield simulation results that are comparable to the erosion measurements. We argue that one possible mechanism for the higher erosion is high energy ions that graze the channel corner of the pole and become trapped in a sheath that develops downstream. Within the sheath, we find that the electric field is large enough to turn the ions towards the surface of the pole. A simplified calculation shows that the erosion rates produced by sputtering of trapped ions are similar to the measurements. We also propose that the presence of local oscillations and high-energy cathode ions may be sources of the enhanced erosion We conclude this article by proposing a combination of analysis, simulation, and experimental measurements that can be used to address the validity of these hypotheses.

Mikellides, Ioannis G.↗

NASA NTRS2011

High Frequency Acoustic Response Characterization and Analysis of the Deep Throttling Common Extensible Cryogenic Engine

The Common Extensive Cryogenic Engine program demonstrated the operation of a deep throttling engine design. The program, spanning five years from August 2005 to July 2010, funded testing through four separate engine demonstration test series. Along with successful completion of multiple objectives, a discrete response of approximately 4000 Hz was discovered and explored throughout the program. The typical low-amplitude acoustic response was evident in the chamber measurement through almost every operating condition; however, at certain off-nominal operating conditions, the response became discrete with higher amplitude. This paper summarizes the data reduction, characterization, and analysis of the 4,000 Hz response for the entire program duration, using the large amount of data collected. Upon first encountering the response, new objectives and instrumentation were incorporated in future test series to specifically collect 4,000 Hz data. The 4,000 Hz response was identified as being related to the first tangential acoustic mode by means of frequency estimation and spatial decomposition. The latter approach showed that the effective node line of the mode was aligned with the manifold propellant inlets with standing waves and quasi-standing waves present at various times. Contour maps that contain instantaneous frequency and amplitude trackings of the response were generated as a significant improvement to historical manual approaches of data reduction presentation. Signal analysis and dynamic data reduction also uncovered several other features of the response including a stable limit cycle, the progressive engagement of subsequent harmonics, the U-shaped time history, an intermittent response near the test-based neutral stability region, other acoustic modes, and indications of modulation with a separate subsynchronous response. Although no engine damage related to the acoustic mode was noted, the peak-to-peak fluctuating pressure amplitude achieved 12.1% of the mean chamber pressure at its highest. The identification of this response in terms of an instability is also discussed.

Casiano, M. J.↗

NASA NTRS2015

Correlation of Amine Swingbed On-Orbit CO2 Performance with a Hardware Independent Predictive Model

The Amine Swingbed Payload is an experimental system deployed on the International Space Station (ISS) that includes a two-bed, vacuum regenerated, amine-based carbon dioxide (CO2) removal subsystem as the principal item under investigation. The aminebased subsystem, also described previously in various publications as CAMRAS 3, was originally designed, fabricated and tested by Hamilton Sundstrand Space Systems International, Inc. (HSSSI) and delivered to NASA in November 2008. The CAMRAS 3 unit was subsequently designed into a flight payload experiment in 2010 and 2011, with flight test integration activities accomplished on-orbit between January 2012 and March 2013. Payload activation was accomplished in May 2013 followed by a 1000 hour experimental period. The experimental nature of the Payload and the interaction with the dynamic ISS environment present unique scientific and engineering challenges, in particular to the verification and validation of the expected Payload CO2 removal performance. A modeling and simulation approach that incorporates principles of chemical reaction engineering has been developed for the amine-based system to predict the dynamic cabin CO2 partial pressure with given inputs of sorbent bed size, process air flow, operating temperature, half-cycle time, CO2 generation rate, cabin volume and the magnitude of vacuum available. Simulation runs using the model to predict ambient CO2 concentrations show good correlation to on-orbit performance measurements and ISS dynamic concentrations for the assumed operating conditions. The dynamic predictive modelling could benefit operational planning to help ensure ISS CO2 concentrations are maintained below prescribed limits and for the Orion vehicle to simulate various operating conditions, scenarios and transients.

Papale, William↗

NASA NTRS2000

Infrared Emission Spectroscopy as a Reliability Tool

In this paper we report on a technique based on infrared emission spectroscopy which has been found to be useful for non-contact measurement of the temperature of a hot spot in the gate channel of a GaAs metal/semiconductor field effect transistor (MESFET). The technique was demonstrated on a powered and un-powered GaAs MESFET attaining a spatial resolution of 0.5 microns. A non-destructive submicron-size spot laser beam provided by an HeNe laser excites an extremely small local area of the gate channel of a GaAs MESFET under various operating conditions. Given the state of the experimental test system, we estimate a spatial resolution of about 0.1 microns and a spectral resolution of about 0. I Angstroms. This provides 15 - 100 times finer spatial resolution than can be obtained using the best passive IR systems available. The temperature resolution (less than 0.02 K/micron) of this technique is depend upon the spectrometer used, and that it can be improved further. The information obtained from this technique can be used to estimate device lifetimes for critical applications and for measurement of channel temperature of devices under actual operating conditions. Another potential use of the novel technique can be as a cost-effective prescreening tool for determining the "hot spot" channel temperature of devices under normal operating conditions which can further improve device design, yield enhancement, and reliable operation

Kim, Q.↗

NASA NTRS1994

Controllability of Free-piston Stirling Engine/linear Alternator Driving a Dynamic Load

This paper presents the dynamic behavior of a Free-Piston Stirling Engine/linear alternator (FPSE/LA) driving a single-phase fractional horse-power induction motor. The controllability and dynamic stability of the system are discussed by means of sensitivity effects of variations in system parameters, engine controller, operating conditions, and mechanical loading on the induction motor. The approach used expands on a combined mechanical and thermodynamic formulation employed in a previous paper. The application of state-space technique and frequency domain analysis enhances understanding of the dynamic interactions. Engine-alternator parametric sensitivity studies, similar to those of the previous paper, are summarized. Detailed discussions are provided for parametric variations which relate to the engine controller and system operating conditions. The results suggest that the controllability of a FPSE-based power system is enhanced by proper operating conditions and built-in controls.

Kankam, M. David↗

NASA NTRS2021

Long-Duration Wear Testing of the ASTRAEUS Thruster, Phase I: 50 kg Xe Throughput

The thruster component of the Ascendant Sub-kW Transcelestial Electric Propulsion System successfully completed a propellant throughput demonstration of 50 kg Xe. The thruster was operated for a total of 3,287 h, of which 97.9% was spent at 300 V – 1000 W, and generated 790 kN-s of total impulse. Thrust, telemetry, and temperature measurements periodically recorded at four operating conditions spanning 200 – 500 V and 200 – 1000 W revealed constant performance across the full test duration. High-fidelity plasma simulations of the erosion of the thruster’s downstream faces suggest a theoretical lifetime capability of >30 kh, corresponding to propellant throughput of >450 kg Xe and a total impulse of >7 MNs, at the 300 V – 1000 W operating condition. The test was voluntarily paused after 50 kg Xe throughput to take these erosion measurements before the test was resumed at a different operating condition.

Goebel, Dan↗

NASA NTRS2013

The Impact of Back-Sputtered Carbon on the Accelerator Grid Wear Rates of the NEXT and NSTAR Ion Thrusters

A study was conducted to quantify the impact of back-sputtered carbon on the downstream accelerator grid erosion rates of the NEXT (NASA's Evolutionary Xenon Thruster) Long Duration Test (LDT1). A similar analysis that was conducted for the NSTAR (NASA's Solar Electric Propulsion Technology Applications Readiness Program) Life Demonstration Test (LDT2) was used as a foundation for the analysis developed herein. A new carbon surface coverage model was developed that accounted for multiple carbon adlayers before complete surface coverage is achieved. The resulting model requires knowledge of more model inputs, so they were conservatively estimated using the results of past thin film sputtering studies and particle reflection predictions. In addition, accelerator current densities across the grid were rigorously determined using an ion optics code to determine accelerator current distributions and an algorithm to determine beam current densities along a grid using downstream measurements. The improved analysis was applied to the NSTAR test results for evaluation. The improved analysis demonstrated that the impact of back-sputtered carbon on pit and groove wear rate for the NSTAR LDT2 was negligible throughout most of eroded grid radius. The improved analysis also predicted the accelerator current density for transition from net erosion to net deposition considerably more accurately than the original analysis. The improved analysis was used to estimate the impact of back-sputtered carbon on the accelerator grid pit and groove wear rate of the NEXT Long Duration Test (LDT1). Unlike the NSTAR analysis, the NEXT analysis was more challenging because the thruster was operated for extended durations at various operating conditions and was unavailable for measurements because the test is ongoing. As a result, the NEXT LDT1 estimates presented herein are considered preliminary until the results of future posttest analyses are incorporated. The worst-case impact of carbon back-sputtering was determined to be the full power operating condition, but the maximum impact of back-sputtered carbon was only a four percent reduction in wear rate. As a result, back-sputtered carbon is estimated to have an insignificant impact on the first failure mode of the NEXT LDT at all operating conditions.

Soulas, George C.↗