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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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Report describes effects of power and temperature overstress on Semtech and Micro Semiconductor devices. Only two Semtech diodes failed catastrophically. Testing on Micro Semiconductor devices stopped because failure limit was reached. Micro diodes suffered lead separation.
External and internal pressure and cold-wall heating-rate distributions were obtained in hypersonic flow on a full-scale heat-sink representation of the space shuttle orbiter wing-elevon-cove configuration in an effort to define effects of flow separation on cove aerothermal environment as a function of cove seal leak area, ramp angle, and free-stream unit Reynolds number. Average free-stream Mach number from all tests was 6.9; average total temperature from all tests was 3360 R; free-stream dynamic pressure ranged from about 2 to 9 psi; and wing angle of attack was 5 deg (flow compression). For transitional and turbulent flow separation, increasing cove leakage progressively increased heating rates in the cove. When ingested mass flow was sufficient to force large reductions in extent of separation, increasing cove leakage reduced heating rates in the cove to those for laminar attached flow. Cove heating-rate distributions calculated with a method that assumed laminar developing channel flow agreed with experimentally obtained distributions within root-mean-square differences that varied between 11 and 36 percent where cove walls were parallel for leak areas of 50 and 100 percent.
Both separate and combined wind tunnel and vibration shaker tests were conducted on two structural panels representative of the Shuttle orbiter in the NASA LaRC 8-foot transonic pressure tunnel (TPT) to determine the effects of combined loads on the thermal protection system (TPS). The primary objective of this test was to provide a combined full-scale load environment and realistic time history of the dynamic pressures, Mach numbers (through transonic), and dynamic structural responses of these panels. The panels were selected from orbiter locations where interactive load sources such as aerodynamic shock waves, turbulent boundary layers, strut-induced vorticity, and substrate deformation combined to provide high bonding loads between the TPS and the orbiter structure. The test panels for this program were highly instrumented with static and dynamic pressure gages, accelerometers, deflectometers, strain gages, Schileren and high-speed photography, and special instrumentation necessary to determine TPS/structure interface loads and tile motions. Two test specimens of each orbiter panel were utilized. Both were high-fidelity representations of the selected orbiter location. The first panel (calibration panel) for each test was utilized to calibrate the tunnel and shaker system as well as provide a data-gathering source for applied static and dynamic loads. The second panel (test panel) for each test was subjected to flight time exposures only so that TPS characteristics as a function of flight mission exposure could be determined. These tests were the last in a series that was suce6ssfully completed and necessary to certify the TPS prior to the first launching of the Shuttle.
Chokshi and Wright presented a detailed model for simulating angular distribution of galaxy images in fields that extended to very high redshifts. Standard tools are used to analyze these simulated galaxy fields for the Omega(O) = 0 and the Omega(O) = 1 cases in order to test the discriminatory power of these tools. Classical number-magnitude diagrams and surface brightness-color-color diagrams are employed to study crowded galaxy fields. An attempt is made to separate the effects due to stellar evolution in galaxies from those due to the space time geometry. The results show that this discrimination is maximized at near-infrared wavelengths where the stellar photospheres are still visible but stellar evolution effects are less severe than those observed at optical wavelenghts. Rapid evolution of the stars on the asymptotic giant branch is easily recognized in the simulated data for both cosmologies and serves to discriminate between the two extreme values of Omega(O). Measurements of total magnitudes of individual galaxies are not essential for studying light distribution in galaxies as a function of redshift. Calculations for the extragalactic background radiation are carried out using the simulated data, and compared to integrals over the evolutionary models used.
A two-equation turbulence model of the K-epsilon type was recently derived by using Renormalization Group (RNG) methods. It was later reported that this RNG based model yields substantially better predictions than the standard K-epsilon model for turbulent flow over a backward facing step - a standard test case used to benchmark the performance of turbulence models in separated flows. The improvements obtained from the RNG K-epsilon model were attributed to the better treatment of near wall turbulence effects. In contrast to these earlier claims, it is shown in this paper that the original version of the RNG K-epsilon model substantially underpredicts the reattachment point in the backstep problem. This is a deficiency that is traced to the modeling of the production of dissipation term. However, with the most recent improvements in the RNG K-epsilon model, excellent results for the backstep problem are now obtained.
A flight investigation of a 1/7-scale rocket-powered model of the XF10F Grumman XFl0F airplane in the swept-wing configuration has been made. The purpose of this test was to determine the static longitudinal stability, damping in pitch, and longitudinal control effectiveness of the airplane with the center of gravity at 20 percent of the wing mean aerodynamic chord. Only a small amount of data was obtained from the test because, immediately after booster separation at a Mach number of 0.88, the configuration was directionally unstable and diverged in sideslip. Simultaneous with the sideslip divergence, the model became longitudinally unstable at 3 degree angle of attack and -6 degree sideslip and diverged in pitch to a high angle of attack. During the pitch-up the free-floating horizontal tail became unstable at 5 degree angle of attack and the tail drifted against its positive deflection limit.
In the present work, the behavior of composite-fastener joints in bearing failure at dynamic stroke rates of 500 in/s, 300 in/s and 100 in/s has been evaluated through progressive damage analysis (PDA) material model in LS-DYNA, namely MAT162. Two joint types: titanium pin and Hi-Lok® fastener were analyzed to identify the differences between without and with preload conditions. A meso-level approach where each lamina was modeled separately was employed and a contact definition based on fracture toughness data was defined to represent composite delamination behavior. Test fixture had been modeled in a detailed manner to account for the dynamic effects and the simulation results were validated against experimental data. Preliminary test-analysis correlation indicated that MAT162 predicted results conservatively when compared to tests. Debris accumulation were observed to greatly affect the test results which were not considered in the current modelling strategies.
The Ares I-X Flight Test Vehicle is the first in a series of flight test vehicles that will take the Ares I Crew Launch Vehicle design from development to operational capability. The test flight is scheduled for April 2009, relatively early in the Ares I design process so that data obtained from the flight can impact the design of Ares I before its Critical Design Review. Because of the short time frame (relative to new launch vehicle development) before the Ares I-X flight, decisions about the flight test vehicle design had to be made in order to complete analysis and testing in time to manufacture the Ares I-X vehicle hardware elements. This paper describes the similarities and differences between the Ares I-X Flight Test Vehicle and the Ares I Crew Launch Vehicle. Areas of comparison include the outer mold line geometry, aerosciences, trajectory, structural modes, flight control architecture, separation sequence, and relevant element differences. Most of the outer mold line differences present between Ares I and Ares I-X are minor and will not have a significant effect on overall vehicle performance. The most significant impacts are related to the geometric differences in Orion Crew Exploration Vehicle at the forward end of the stack. These physical differences will cause differences in the flow physics in these areas. Even with these differences, the Ares I-X flight test is poised to meet all five primary objectives and six secondary objectives. Knowledge of what the Ares I-X flight test will provide in similitude to Ares I as well as what the test will not provide is important in the continued execution of the Ares I-X mission leading to its flight and the continued design and development of Ares I.
The Ares I-X Flight Test Vehicle is the first in a series of flight test vehicles that will take the Ares I Crew Launch Vehicle design from development to operational capability. Ares I-X is scheduled for a 2009 flight date, early enough in the Ares I design and development process so that data obtained from the flight can impact the design of Ares I before its Critical Design Review. Decisions on Ares I-X scope, flight test objectives, and FTV fidelity were made prior to the Ares I systems requirements being baselined. This was necessary in order to achieve a development flight test to impact the Ares I design. Differences between the Ares I-X and the Ares I configurations are artifacts of formulating this experimental project at an early stage and the natural maturation of the Ares I design process. This paper describes the similarities and differences between the Ares I-X Flight Test Vehicle and the Ares I Crew Launch Vehicle. Areas of comparison include the outer mold line geometry, aerosciences, trajectory, structural modes, flight control architecture, separation sequence, and relevant element differences. Most of the outer mold line differences present between Ares I and Ares I-X are minor and will not have a significant effect on overall vehicle performance. The most significant impacts are related to the geometric differences in Orion Crew Exploration Vehicle at the forward end of the stack. These physical differences will cause differences in the flow physics in these areas. Even with these differences, the Ares I-X flight test is poised to meet all five primary objectives and six secondary objectives. Knowledge of what the Ares I-X flight test will provide in similitude to Ares I - as well as what the test will not provide - is important in the continued execution of the Ares I-X mission leading to its flight and the continued design and development of Ares I.
An investigation of the aerodynamic performance of leading-edge flaps on three clipped delta and three clipped double-delta wing planforms with aspect ratios of 1.75, 2.11, and 2.50 was conducted in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.60, 1.90, and 2.16. A primary set of fullspan leading-edge flaps with similar root and tip chords were investigated on each wing, and several alternate flap planforms were investigated on the aspect-ratio-1.75 wings. All leading-edge flap geometries were effective in reducing the drag at lifting conditions over the range of wing aspect ratios and Mach numbers tested. Application of a primary flap resulted in better flap performance with the double-delta planform than with the delta planform. The primary flap geometry generally yielded better performance than the alternate flap geometries tested. Trim drag due to flap-induced pitching moments was found to reduce the leading-edge flap performance more for the delta planform than for the double-delta planform. Flow-visualization techniques showed that leading-edge flap deflection reduces crossflow shock-induced separation effects. Finally, it was found that modified linear theory consistently predicts only the effects of leading-edge flap deflection as related to pitching moment and lift trends.
Regulations to establish operational and performance requirements for unmanned aircraft systems (UAS) are being developed by a consortium of government, industry and academic institutions (RTCA, 2013). Those requirements will apply to the new detect-and-avoid (DAA) systems and other equipment necessary to integrate UAS with the United States (U.S) National Airspace System (NAS) and will be determined according to their contribution to the overall safety case. That safety case requires demonstration that DAA-equipped UAS collectively operating in the NAS meet an airspace safety threshold (AST). Several key gaps must be closed in order to link equipment requirements to an airspace safety case. Foremost among these is calculation of the systems risk ratio, the degree to which a particular system mitigates violation of an aircraft separation standard (FAA, 2013). The risk ratio of a DAA system, in combination with risk ratios of other collision mitigation mechanisms, will determine the overall safety of the airspace measured in terms of the number of collisions per flight hour. It is not known what the effectiveness is of a pilot-in-the-loop DAA system or even what parameters of the DAA system most improve the pilots ability to maintain separation. The relationship between the DAA system design and the overall effectiveness of the DAA system that includes the pilot, expressed as a risk ratio, must be determined before DAA operational and performance requirements can be finalized. Much research has been devoted to integrating UAS into non-segregated airspace (Dalamagkidis, 2009, Ostwald, 2007, Gillian, 2012, Hesselink, 2011, Santiago, 2015, Rorie 2015 and 2016). Several traffic displays intended for use as part of a DAA system have gone through human-in-the-loop simulation and flight-testing. Most of these evaluations were part of development programs to produce a deployable system, so it is unclear how to generalize particular aspects of those designs to general requirements for future traffic displays (Calhoun, 2014). Other displays have undergone testing to collect data that may generalize to new displays, but have not been evaluated in the context of the development of an overall safety case for UAS equipped with DAA systems in the NAS (Bell, 2012). Other research efforts focus on DAA surveillance performance and separation standards. Together with this work, they are expected to facilitate validation of the airspace safety case (Park, 2014 and Johnson, 2015). The contribution of the present work is to quantify the effectiveness of the pilot-automation system to remain well clear as a function of display features and surveillance sensor error. This quantification will help enable selection of a minimum set of DAA design features that meets the AST, a set that may not be unique for all UAS platforms. A second objective is to collect and analyze pilot performance parameters that will improve the modeling of overall DAA system performance in non-human-in-the-loop simulations. Simulating the DAA-equipped UAS in such batch experiments will allow investigation of a much larger number of encounters than is possible in human simulations. This capability is necessary to demonstrate that a particular set of DAA requirements meets the AST under all foreseeable operational conditions.
Experimental verification of the theory describing arbitrary motions of an airfoil is reported. The experimental apparatus is described. A mechanism was designed to provide two separate degrees of freedom without friction or backlash to mask the small but important aerodynamic effects of interest.
In this paper a pair of observables are proposed as alternative ways, by examining the fluctuation of net momentum-ordering of charged pairs, to study the charge separation induced by the Chiral Magnetic Effect (CME) in relativistic heavy ion collisions. They are, the out-of-plane to in-plane ratio of fluctuation of the difference between signed balance functions measured in pair’s rest frame, and the ratio of it to similar measurement made in the laboratory frame. In this work, both observables have been studied with simulations including flow-related backgrounds, and for the first time, backgrounds that are related to resonance's global spin alignment. The two observables have similar positive responses to signal, and opposite, limited responses to identifiable backgrounds arising from resonance flow and spin alignment. Both observables have also been tested with two realistic models, namely, a multi-phase transport (AMPT) model and the anomalous-viscous fluid dynamics (AVFD) model. These two observables, when cross examined, will provide useful insights in the study of CME-induced charge separation.
This paper reports results from pressure tests on an ogive-cylinder in the low-turbulence 12-foot pressure wind tunnel. The results consist of pressure distributions over a wide range of Reynolds numbers and angles of attack. The tests encompassed a complete coverage of different roll orientations shown to be essential in order to fully define all the possible flow conditions. When the various roll-angle results are combined, it is possible to interpret the effects of changing angle of attack or Reynolds number. Two basic mechanisms for producing asymmetric flow are identified. One mechanism operates in both the laminar and the fully turbulent separation regimes; this mechanism is the one qualitatively described by the impulsive flow analogy. The other mechanism occurs only in the transitional separation regime. This asymmetric flow has the same form as that in the two-dimensional crossflow on a circular cylinder in the transitional flow regime. Finally, these results make it possible to draw up critical Reynolds number boundaries between the laminar, transitional, and fully turbulent separation regimes throughout the angle-of-attack range from 20 to 90 deg.
The aerodynamic database for the Orion Launch Abort System (LAS) was developed largely from wind tunnel tests involving powered jet simulations of the rocket exhaust plumes, supported by computational fluid dynamics (CFD) simulations. The LAS contains three solid rocket motors used in various phases of an abort to provide propulsion, steering, and Launch Abort Tower (LAT) jettison from the Crew Module (CM). This paper describes a pair of wind tunnel experiments performed at transonic and supersonic speeds to determine the aerodynamic effects due to proximity and jet interactions during LAT jettison from the CM at the end of an abort. The tests were run using two different scale models at angles of attack from 150deg to 200deg , sideslip angles from -10deg to +10deg , and a range of powered thrust levels from the jettison motors to match various jet simulation parameters with flight values. Separation movements between the CM and LAT included axial and vertical translations as well as relative pitch angle between the two bodies. The paper details aspects of the model design, nozzle scaling methodology, instrumentation, testing procedures, and data reduction. Sample data are shown to highlight trends seen in the results.
A description of the Rotor Systems Research Aircraft (RSRA) was presented, with particular emphasis on the unique systems that provide the potential for good flight/ground test facility correlation. These flight research vehicles are designed specifically with the capabilities necessary for the effective and efficient in-flight test and verification of promising new rotor concepts and supporting technology developments. The research mission and unique features of the RSRA will provide the capability to measure and separate the flight loads of the airframe and rotor, thereby allowing direct flight/ground test facility correlation of rotors as well as correlation with analytical models.
A dynamic model used for verification of analytical and experimental methods is documented. The model consists of five substructures to simulate the multibody, low frequency nature of large space structures. Design considerations which led to a fundamental vibration frequency of less than one Hz are described. Finite element analysis used to predict the vibration modes and frequencies of the experimental model is presented. In addition, modeling of cable suspension effects using prestressed vibration analysis is described. Details of the expermental and analytical models are included to permit replication of the study. Results of the modal vibration tests and analysis are presented in a separate document.
Dynamic measurements of fluctuating static pressure levels were made using flush mounted high frequency response pressure transducers at eleven locations in the circuit of the National Transonic Facility (NTF) over the complete operating range of this wind tunnel. Measurements were made at test section Mach numbers from 0.2 to 1.2, at pressure from 1 to 8.6 atmospheres and at temperatures from ambient to -250 F, resulting in dynamic flow disturbance measurements at the highest Reynolds numbers available in a transonic ground test facility. Tests were also made independently at variable Mach number, variable Reynolds number, and variable drivepower, each time keeping the other two variables constant thus allowing for the first time, a distinct separation of these three important variables. A description of the NTF emphasizing its flow quality features, details on the calibration of the instrumentation, results of measurements with the test section slots covered, downstream choke, effects of liquid nitrogen injection and gaseous nitrogen venting, comparisons between air and nitrogen, isolation of the effects of Mach number, Reynolds number, and fan drive power, and identification of the sources of significant flow disturbances is included. The results indicate that primary sources of flow disturbance in the NTF may be edge-tones generated by test section sidewall re-entry flaps and the venting of nitrogen gas from the return leg of the tunnel circuit between turns 3 and 4 in the cryogenic mode of operation. The tests to isolate the effects of Mach number, Reynolds number, and drive power indicate that Mach number effects predominate. A comparison with other transonic wind tunnels shows that the NTF has low levels of test section fluctuating static pressure especially in the high subsonic Mach number range from 0.7 to 0.9.