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Apparatus, techniques, and instrumentation for experimental correlation of nonreactive injection schemes and combustion effects in liquid propellant rocket engine
This paper addresses the regulatory processes and requirements already in place by which an applicant might obtain experimental airworthiness certification for a civil Unmanned Aircraft System (UAS). It is more extensive and subsequent to an earlier, similar deliverable, PD007, which was an interim study of the same topic. Since few regulatory airworthiness and operating standards exist for UAS like those for traditional manned aircraft and since most UAS have historically been developed and operated under military auspices, civil use of UAS in the NAS is a new and unfamiliar challenge requiring specific and unique considerations. Experimental certification is the most basic level of FAA approval toward routine UAS operation in the NAS. The paper reviews and explains existing FAA requirements for an applicant seeking experimental airworthiness approval and details the process for submission of necessary information. It summarizes the limited purposes for which experimental aircraft may be used and addresses pertinent aspects of UAS design, construction and operation in the NAS in harmony with traditional manned aircraft. Policy IPT position is that UAS, while different from manned aircraft, can use the same initial processes to gain civil operating experience under the experimental approval. Particular note is taken of those UAS-unique characteristics which require extra attention to assure equivalent safety of operation, such as the UAS control station and sense-and-avoid. The paper also provides "best practices" guidance for UAS manufacturers and FAA personnel in two appendices. The material in Appendix A is intended to provide guidance on assuring UAS safety to FAA, and provides FAA personnel with a suggested list of items to review, with a focus on UAS unique factors, prior to issuance of an experimental airworthiness certificate. Appendix B provides an outline for a program letter which a manufacturer could use in preparing the application for an UAS experimental airworthiness certificate.
This Review Article compares new experimental and corresponding semiclassical Stark broadened data for non-hydrogenic spectral lines of neutral atoms and positive ions. This Review covers the period 2008 until the end of 2020 and presents the continuation of previous critical reviews from 1976 (two), 1984 (two), 1990, 2002, and one from 2009. This Review reports the Stark broadening parameters of 1665 spectral lines belonging to 35 elements with 61 different atomic and ion species. The data are taken from 60 papers. Experimental data are arranged by elements and spectra, and these are reported in tables in alphabetical and numerical order, respectively. Each experimental Stark broadening parameter is followed with estimated accuracy. The experimental Stark broadening and shift data presented in the tables are compared with the corresponding semiclassical results available in the literature. In addition, for comparison experiments versus semiclassical theoretical Stark broadening parameters, a numerically improved theoretical approach developed based on Griem and co-workers’ theory is used to evaluate Stark widths and shifts of all studied lines whenever required atomic energy levels data and transition probabilities are available. At the end of the text report, for each analyzed neutral or ion species, the information about the location of the same species data in preceding reviews is given.
The results of a three-year experimental campaign aimed at comprehensively documenting the separated flow over a three-dimensional bump are presented with the purpose of generating a benchmark experimental database useful in validating computational fluid dynamics (CFD) flow simulations. The bump model geometry was designed to provide well-defined and repeatable smooth-body flow separation conditions that were suitable for both experiments and simulations. The bump had a Gaussian streamwise profile with a constant height equal to 8.5% of its width over the central 60% of the test section width. The remaining 40% were outboard spanwise portions that gradually taper to zero using an error function profile to minimize side-wall boundary layer effects. The model was immersed in a turbulent boundary layer that was developed on a suspended flat plate in the Notre Dame Mach 0.6 Wind Tunnel. In order to document the effect of the incoming boundary layer thickness on the flow separation, the bump model could be located at two streamwise positions. The mean velocity and turbulence intensity of the wind tunnel freestream flow field and approaching turbulent boundary layer were fully documented. The measurements of the flow separation region included surface visualization, wall shear stress using oil-film interferometry, mean and dynamic surface pressure, and planar and stereoscopic particle image velocimetry. The experiments were conducted over a range of Mach numbers from 0.05 to 0.2 corresponding to a range of Reynolds numbers based on the test section spanwise dimension (0.914m) of 1.0 × 10 6 ≤ Re L = U ∞ L/ν ≤ 4.0 × 10 6 . The bulk of the results are presented for the higher Mach number conditions of 0.1 and 0.2 with Re L = 2.0 × 10 6 and 4.0 × 10 6 , respectively. Extensive uncertainty analysis of the data was performed. The data is archived in the NASA Langley Turbulence Modeling Resource website at https://turbmodels.larc.nasa.gov/Other_exp_Data/speedbump_sep_exp.html. In addition to the experiments, a computational effort was made in parallel by the CFD group at Boeing Research & Technology highlighting the usefulness of the data set, which is outlined in the accompanying CFD report.
A lifting airfoil theoretically designed for shockless supercritical flow utilizing a complex hodograph method has been evaluated in the Langley 8-foot transonic pressure tunnel at design and off-design conditions. The experimental results are presented and compared with those of an experimentally designed supercritical airfoil which were obtained in the same tunnel.
Wind-tunnel tests on the British Experimental Rotor Programme (BERP) tip are described, and the results are compared with computational fluid dynamics (CFD) results. The test model was molded using the Lynx-BERP blade tooling to provide a semispan, cantilever wing comprising the outboard 30 percent of the rotor blade. The tests included both surface-pressure measurements and flow visualization to obtain detailed information of the flow over the BERP tip for a range of angles of attack. It was observed that, outboard of the notch, favorable pressure gradients exist which ensure attached flow, and that the tip vortex also remains stable to large angles of attack. On the rotor, these features yield a very gradual break in control loads when the retreating-blade limit is eventually reached. Computational and experimental results were generally found to be in good agreement.
Experimental profiles for heat pipe startup from the frozen state were obtained, using a high-temperature sodium/stainless steel pipe with multiple heat sources and sinks to investigate the startup behavior of the heat pipe for various heat loads and input locations, with both low and high heat rejection rates at the condensor. The experimental results of the performance characteristics for the continuum transient and steady-state operation of the heat pipe were analyzed, and the performance limits for operation with varying heat fluxes and location are determined.
NASA grant NAG9-329 was in effect from 3/1/89 to 8/31/94, the last 18 months being a no-cost extension. While the grant was in effect, the P.I., coworkers, and students gave 22 talks and poster sessions at professional meetings, published 12 articles in referred journals (one more is in press, and another is in review), and edited 2 workshop reports relevant to this project. Copies of all the publications are appended to this report. The major accomplishments during the grant period have derived from three quarters: 1) the application of quantitative models of fractional crystallization and partial melting to various problems in planetary science, such as the petrogenesis of picritic glasses and mare basalts and the implications of the SNC meteorites for martian evolution; 2) an experimental study of silicate liquid immiscibility relevant to early lunar differentiation and the petrogenesis of evolved highlands rocks; and 3) experimental studies of massif anorthosites and related rocks that provide terrestrial analogs for the proposed origin of lunar anorthosites by multistage processes. The low-pressure aspects of the quantitative models were developed by the P.I. in the 1980s with NASA support and culminated with a paper comparing the crystallization of terrestrial and lunar lavas. The basis for the high-pressure modifications to the quantitative models is a data set gleaned from high pressure melting experiments done at Lamont and is supplemented by published data from other labs that constrain the baric and compositional dependences of various liquidus phase boundaries such as olivine/orthopyroxene, relevant to the melting of the mantles of the terrestrial planets. With these models it is possible to predict not only the thermal and compositional evolution of magmatic liquids ranging in composition from lumar mare basalt to terrestrial calc-alkaline basalts, but also the small increments of fractional melting that are produced when mantle rises adiabatically. Copies of the crystallization/melting programs have been given to several colleagues in planetary science. Additionally, a series of computer graphics programs, based on the algorithms in the crystallization programs have been developed that display liquidus diagrams appropriate to input compositions.
A combined experimental and computational study has been performed to investigate the detailed heat transfer coefficient distributions within a complex blade trailing edge passage. The experimental measurements are made using a steady liquid crystal thermography technique applied to one major side of the passage. The geometry of the trailing edge passage is that of a two-pass serpentine circuit with a sharp 180-degree turning region at the tip. The upflow channel is split by interrupted ribs into two major subchannels, one of which is turbulated. This channel has an average aspect ratio of roughly 14:1. The spanwise extent of the channel geometry includes both area convergence from root to tip, as well as taper towards the trailing edge apex. The average section Reynolds numbers tested in this upflow channel range from 55,000 to 98,000. The tip section contains a turning vane near the extreme comer. The downflow channel has an aspect ratio of about 5:1, and also includes convergence and taper. Turbulators of varying sizes are included in this channel also. Both detailed heat transfer and pressure distribution measurements are presented. The pressure measurements are incorporated into a flow network model illustrating the major loss contributors.
Results are presented of a cold-flow and hot-fire experimental study of the mixing and atomization characteristics of injector elements incorporating noncircular orifices. Both liquid/liquid and gas/liquid element types are discussed. Unlike doublet and triplet elements (circular orifices only) were investigated for the liquid/liquid case while concentric tube elements were investigated for the gas/liquid case. It is concluded that noncircular shape can be employed to significant advantage in injector design for liquid rocket engines.
Tests were conducted on a linear cascade of airfoils oscillating in pitch to measure the unsteady pressure response on selected blades along the leading edge plane of the cascade, over the chord of the center blade, and on the sidewall in the plane of the leading edge. The pressure data were reduced to Fourier coefficient form for direct comparison, and were also processed to yield integrated loads and, particularly, the aerodynamic damping coefficient. Results from the unsteady Verdon/Caspar theory for cascaded blades with nonzero thickness and camber were compared with the experimental measurements. The three primary results are: (1) from the leading edge plane blade data, the cascade was judged to be periodic in unsteady flow over the range of parameters tested; (2) the interblade phase angle was found to be the single most important parameter affecting the stability of the oscillating cascade blades; and (3) the real blade theory and the experiment were in excellent agreement for the several cases chosen for comparison.
An experimental low speed study of the separating confluent boundary layer on a NASA GAW-1 high lift airfoil is described. The airfoil was tested in a variety of high lift configurations comprised of leading edge slat and trailing edge flap combinations. The primary test instrumentation was a two dimensional laser velocimeter (LV) system operating in a backscatter mode. Surface pressures and corresponding LV derived boundary layer profiles are given in terms of velocity components, turbulence intensities and Reynolds shear stresses as characterizing confluent boundary layer behavior up to and beyond stall. LV derived profiles and associated boundary layer parameters and those obtained from more conventional instrumentation such as pitot static transverse, Preston tube measurements and hot-wire surveys are compared.
Results of experimental electric propulsion research are presented. A plasma generator, with an arc stabilized by an air vortex is examined. The heat transfer efficiency between arc and fluid environment at a varying current and flow rate is discussed.
Earlier results from computer simulation studies suggest a correlation between the spatial distribution of stacking errors in the Beta-SiC structure and features observed in X-ray diffraction patterns of the material. Reported here are experimental results obtained from two types of nominally Beta-SiC specimens, which yield distinct XRD data. These samples were analyzed using high resolution transmission electron microscopy (HRTEM) and the stacking error distribution was directly determined. The HRTEM results compare well to those deduced by matching the XRD data with simulated spectra, confirming the hypothesis that the XRD data is indicative not only of the presence and density of stacking errors, but also that it can yield information regarding their distribution. In addition, the stacking error population in both specimens is related to their synthesis conditions and it appears that it is similar to the relation developed by others to explain the formation of the corresponding polytypes.
Summary of preliminary experimental results and observations from a subscale, 2D, inert gas supersonic plume-surface interaction test campaign.
High altitude base heating of 4-engine rocket nozzle configuration
Times of consciousness, collapse, and survival and pathology examinations of dogs rapidly decompressed to near vacuum environment
Test equipment which was built and the calibration tests are described. The test equipment consists of a two-bladed rotor of 16-in. diameter, the blades of which are elastically hinged in flapping. The feathering shaft of the blades can be harmonically rotated with the help of a cam mechanism located inside the hollow rotor shaft. The frequency range measured in the rotating system can be adjusted between 0 and 80 cps and the rotor speed between 0 and 40 cps. The test equipment is for measuring the flapping response of the blades to harmonic feathering excitation.