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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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A fan-driven transonic cryogenic tunnel was designed, and its purging, cooldown, and warmup times were determined satisfactory. Cooling with liquid nitrogen is at the power levels required for transonic testing. Good temperature distributions are obtained by using a simple nitrogen injection system.
Male Swiss albino mice were exposed to the pyrolysis products of two fire-retardant materials, a chlorinated aromatic polyamide and a copolymer of vinylidine fluoride and hexafluoropropene. Comparison tests were made with cotton and a 50/50 cotton-polyester composite. In addition, tests were conducted under the presence of CO, and mice were injected intraperitoneally or intramuscularly with aqueous solutions containing dissolved effluents from the pyrolysis of cotton or of chlorinated aromatic polyamide. Results indicate that unique thermodecomposition products of the polymeric materials are more toxic to mice than are other products from cotton under similar controlled conditions.
In liquid injection thrust vector control, a rocket jet is deflected for steering purposed by injecting a liquid into the nozzle exit cone. The liquid is preferably both dense and reactive so that it adds mass and energy and generates shocks in the supersonic exhaust. This behavior increases thrust in the affected part of the jet producing not only a side force for steering but an addition to axial thrust. This paper presents a summary of current liquid injection thrust vector control technology, including procedures for design, development, analysis, testing and evaluation, together with supporting data and references.
Laser modeling activity is described which involved addition of an option allowing N2 as a second diatomic gas. This option is now operational and a few test cases involving N2/CO mixtures were run. Results from these initial test cases are summarized. In the laboratory, a CW double-discharge test facility was constructed and tested. Features include: water-cooled removable electrodes, O-ring construction to facilitate cleaning and design modifications, increased discharge length, and addition of a post-discharge observation section. Preliminary tests with this facility using N2 yielded higher power loadings than obtained in the first-generation facility. Another test-section modification, recently made and as yet untested, will permit injection of secondary gases into the cathode boundary layer. The objective will be to vary and enhance the UV emission spectrum from the auxiliary discharge, thereby influencing the level of photoionization in the main discharge region.
A theoretical and experimental study has been made of the effect of wing-mounted fins on the vortex wakes of subsonic aircraft. The theory is used (a) to gain an understanding of wake alleviation by vortex injection and (b) to guide the experimental investigation. Wind-tunnel tests were used to evaluate the alleviation achievable and to find the optimum values for the various fin parameters. It was found that vertical fins mounted on the upper surface of a wing could lower the wake-induced rolling moments on an encountering wing by a factor of 3 or more. The most promising fin configuration found for the Boeing 747 model is a fin positioned 48% outboard from the centerline to the wingtip with a height of 0.014 span, a chord of 0.085 span, and an 18 deg angle of attack. This fin configuration caused a 10% increase in drag but no lift penalty.
The results of a recent wind tunnel test program to evaluate the effectiveness of the Tip Air Mass Injection (TAMI) system in modifying the blade tip vortex occurring during helicopter flight is described with attention to the effect of this modification on the impulsive noise. The measurement program is explained, and the correlation between experimental and predicted results is discussed. Topics considered include the effect of descent rate on noise pressure time histories, the effect of air mass injection on noise, and the analysis based on a dB(A) weighted approach. Impulsive noise generated by the interaction of a helicopter rotor blade and the concentrated tip vortex during forward flight descent is a primary contributor to acoustic annoyance as it draws early attention to the presence of the helicopter.
The condensation research at NASA Langley Research Center has used a variety of experimental approaches to gather information on seed particles that can act as sites for condensation growth. Total pressure measurements have suggested that condensation growth is caused by impurities in the flow and not by unevaporated liquid nitrogen (LN2) droplets resulting from the LN2 injected to cool the 0.3-Meter Transonic Cryogenic Tunnel. A separate test with an optical droplet sizing probe, which was designed to detect droplets in the 2- to 300-micron range, confirmed the conclusions from the total pressure measurements and also discovered what appears to be solidified oil droplets having diameters of about 3 microns. These oil droplets appear to be the dominant source of seed particles above 2 microns. However, computer simulations of static pressure test data suggest that the measured condensation effects are the result of more numerous, smaller seeds with number densities on the order of 10 to the 12th per kilogram of the gas and diameters on the order of 0.5 microns.
The results of a test-particle simulation studying the movement of a lithium tracer ion injected upstream of the bow shock are reported. Wave activity consists of parallel and antiparallel propagating Alfven waves characterized by a frequency power spectrum within a frequency or range of amplitudes defined separately in the upstream and downstream regions. The results show that even a moderate level of wave activity can substantially change the results obtained in the absence of waves. Among the effects observed are: (1) increased ion transmission; (2) both the average energy gain and spread about the average are increased for transmitted and reflected particles; (3) the average final pitch angle for transmitted particles tends to 90 deg, and the spread of reflected particles is reduced; and (4) the spatial dispersion of the ions on the bow shock after a single encounter is increased.
A collection of document abstracts are presented on the topic of improving software dependability through NASA grant NAG-1-1123. Specific topics include: modeling of error detection; software inspection; test cases; Magnetic Stereotaxis System safety specifications and fault trees; and injection of synthetic faults into software.
A variety of different types of actuators have been previously investigated as flow control devices. Potential applications include the control of boundary layer separation in external flows, as well as jet engine inlet and diffuser flow control. The operating principles for such devices are typically based on either mechanical deflection of control surfaces (which include MEMS flap devices), mass injection (which includes combustion driven jet actuators), or through the use of synthetic jets (diaphragm devices which produce a pulsating jet with no net mass flow). This paper introduces some of the initial flow visualization work related to the development of a relatively new type of combustion-driven jet actuator that has been proposed based on a pulse detonation principle. The device is designed to utilize localized detonation of a premixed fuel (Hydrogen)-air mixture to periodically inject a jet of gas transversely into the primary flow. Initial testing with airflow successfully demonstrated resonant conditions within the range of acoustic frequencies expected for the design. Schlieren visualization of the pulsating air jet structure revealed axially symmetric vortex flow, along with the formation of shocks. Flow visualization of the first successful sustained oscillation condition is also demonstrated for one configuration of the current test section. Future testing will explore in more detail the onset of resonant combustion and the approach to conditions of sustained resonant detonation.
An investigation is performed to evaluate the performance of a computational fluid dynamics (CFD) tool for the prediction of the reacting flow in a liquid-fueled combustor that uses water injection for control of pollutant emissions. The experiment consists of a multisector, liquid-fueled combustor rig operated at different inlet pressures and temperatures, and over a range of fuel/air and water/fuel ratios. Fuel can be injected directly into the main combustion airstream and into the cavities. Test rig performance is characterized by combustor exit quantities such as temperature and emissions measurements using rakes and overall pressure drop from upstream plenum to combustor exit. Visualization of the flame is performed using gray scale and color still photographs and high-frame-rate videos. CFD simulations are performed utilizing a methodology that includes computer-aided design (CAD) solid modeling of the geometry, parallel processing over networked computers, and graphical and quantitative post-processing. Physical models include liquid fuel droplet dynamics and evaporation, with combustion modeled using a hybrid finite-rate chemistry model developed for Jet-A fuel. CFD and experimental results are compared for cases with cavity-only fueling, while numerical studies of cavity and main fueling was also performed. Predicted and measured trends in combustor exit temperature, CO and NOx are in general agreement at the different water/fuel loading rates, although quantitative differences exist between the predictions and measurements.
Stiff yet ultra-light lattice structures constructed using digital materials have many practical applications as the building block for aircraft and other structures. By furthering our understanding of how material configuration affects the structural properties of an ultralight lattice, we can intelligently design these structures based on their intended function. Here we compare the behavior of ultralight lattice structures when fabricated by different materials. The individual unit cells of the lattice structures are referred to as voxels. The stiffness, elastic modulus, and yield strength of the specimens in compression and tension are determined through mechanical testing. Specimens are tested both as single voxel as well as 4x4x4 voxel constructions on an Instron 5982 Universal Testing System until failure. Each voxel is manufactured in bulk through injection molding, with a unit cell pitch of 76.2 mm. Individual voxels are fastened with machine screws and nuts to create assemblies. Four separate materials are used as voxel compositions in this experiment. These include a homogeneous polymer referred to as Ultem 1000, a glass-fiber reinforced polymer referred to as Ultem 2200, a polymer with chopped carbon fibers as 30% of its fill, and homogenous polypropylene. This work compares mechanical behavior, as well as the convergence behavior of the lattice as the size of the lattice assembly increases for various materials. The goal of this study is to characterize the behavior of homogenous lattices such that heterogenous lattices can be designed with different material voxels to achieve target material properties for ultralight space applications.
This paper presents experimental results for a low-NOxaero gas turbine combustor, in particular, a third-generationswirl-venturi lean direct injection (SV-LDI-3) combustor conceptcalled V4. The purpose of testing was three-fold. First,to evaluate the combustor against the 80% NOx reduction goalset by NASA’s AATT project. Second, to compare V4 to a previousSV-LDI-3 combustor concept called V3, especially at lowpower conditions. Third, to examine the accuracy of a type ofcorrelation equation frequently used by engine systems analysisgroups to estimate NOx emissions. All three testing goals weremet. For the first testing goal, with an estimated NOx reductionof 85%-90%, SV-LDI-3 V4 surpassed the AATT goal. For thesecond goal, however, V4 did not perform better than V3 at lowpower conditions. For the third goal, it was found that a majorassumption of the correlation equations — a simple dependenceon combustor inlet pressure — did not hold.
A simulated lightning test was conducted on the backup spacecraft for the Apollo Soyuz Test Project mission (CSM-119) to determine the susceptibility of the Apollo spacecraft to damage from the indirect effects of lightning. It is demonstrated that induced lightning effects from low-level injected currents can be scaled linearly to those which are obtained in a full threat lightning stroke. Test results indicate that: (1) many of the power and signal critical circuits would fail if subjected to full-threat lightning, (2) pyrotechnic circuits are safe for full-threat lightning, and (3) common-mode voltages exceeded the failure criteria level for all but three of the circuits tested.
Combustion at inlet-air temperatures of 1100 to 1250 K was studied for application to advanced automotive gas turbine engines. Combustion was initiated by the hot environment, and therefore no external ignition source was used. Combustion was stabilized without a flameholder. The tests were performed in a 12-cm-diameter test section at a pressure of 250,000 Pa, with reference velocities of 32 to 60 m/s and at maximum combustion temperatures of 1350 to 1850 K. Number 2 diesel fuel was injected by means of a multiple-source fuel injector. Unburned hydrocarbon emissions were negligible for all test conditions. Nitrogen oxide emissions were less than 1.9 g NO2/kg fuel for combustion temperatures below 1680 K. Carbon monoxide emissions were less than 16 g CO/kg fuel for combustion temperatures greater than 1600 K, inlet-air temperatures higher than 1150 K, and residence times greater than 4.3 ms.
Development tests are being conducted to characterize unconfined Hydrogen/air and Hydrogen/Oxygen blast characteristics. Most of the existing experiments for these types of explosions address contained explosions, like shock tubes. Therefore, the Hydrogen Unconfined Combustion Test Apparatus (HUCTA) has been developed as a gaseous combustion test device for determining the relationship between overpressure, impulse, and flame speed at various mixture ratios for unconfined reactions of hydrogen/oxygen and hydrogen/air. The system consists of a central platform plumbed to inject and mix component gasses into an attached translucent bag or balloon while monitoring hydrogen concentration. All tests are ignited with a spark with plans to introduce higher energy ignition sources in the future. Surrounding the platform are 9 blast pressure "Pencil" probes. Two high-speed cameras are used to observe flame speed within the combustion zone. The entire system is raised approx. 6 feet off the ground to remove any ground reflection from the measurements. As of this writing greater than 175 tests have been performed and include Design of Experiments test sets. Many of these early tests have used bags or balloons between approx. 340L and approx. 1850L to quantify the effect of gaseous mixture ratio on the properties of interest. All data acquisition is synchronized between the high-speed cameras, the probes, and the ignition system to observe flame and shock propagation. Successful attempts have been made to couple the pressure profile with the progress of the flame front within the combustion zone by placing a probe within the bag. Overpressure and impulse data obtained from these tests are used to anchor engineering analysis tools, CFD models and in the development of blast and fragment acceleration models.
A program was conducted to develop the technology for joints and cutouts in a composite fuselage that meets all design requirements of a large transport aircraft for the 1990s. An advanced trijet derivative of the DC-10 was selected as the baseline aircraft. Design and analysis of a 30-foot-long composite fuselage barrel provided a realistic basis for the test effort. The primary composite material was Hexcel F584 resin on 12 K IM6 fiber, in tape and broadgoods form. Fiberglass broadgoods were used in E-glass and S-glass fiber form in the cutout region of some panels. Additionally, injection-molded chopped graphite fiber/PEEK was used for longeron-to-frame shear clips. The test effort included four groups of test specimens, beginning with coupon specimens of mono-layer and cross-piled laminates, progressing through increasingly larger and more complex specimens, and ending with two 4- by 5-foot curved fuselage side panels. One of the side panels incorporated a transverse skin splice, while the second included two cabin window cutouts.
Combustion at inlet air temperatures of 1100 to 1250 K was studied for application to advanced automotive gas turbine engines. Combustion was initiated by the hot environment, and therefore no external ignition source was used. Combustion was stabilized without a flameholder. The tests were performed in a 12 cm diameter test section at a pressure of 2.5 x 10 to the 5th power Pa, with reference velocities of 32 to 60 m/sec and at maximum combustion temperatures of 1350 to 1850 K. Number 2 diesel fuel was injected by means of a multiple source fuel injector. Unburned hydrocarbons emissions were negligible for all test conditions. Nitrogen oxides emissions were less than 1.9 g NO2/kg fuel for combustion temperatures below 1680 K. Carbon monoxide emissions were less than 16 g CO/kg fuel for combustion temperatures greater than 1600 K, inlet air temperatures higher than 1150 K, and residence times greater than 4.3 microseconds.