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Sonic boom wave pressure history prediction on arbitrarily oriented plane walls by explicit fixed mesh time dependent numerical method
Quantitative gas liquid chromatography analysis of amino acids in biological materials, discussing ion exchange techniques
Analyses of biological substances and geochemical samples, of both terrestrial and extraterrestrial origin, are reported for amino acids at the 1-10 nanogram per gram level achieved by gas-liquid and ion-exchange chromatographic methods. These studies have shown that nanogram quantities of amino acids, present in water extracts of geochemical samples or in other samples of low amino acid concentration, can be successfully determined only if the researcher is well aware of the possible sources of contamination.
The study was primarily directed toward the examination of Apollo 14 lunar fines for indigenous amino acids or materials which could be converted to amino acids on hydrolysis with 6 N hydrochloric acid. Initial experiments were conducted to confirm the integrity of the derivatization reactions and reagents, and to optimize the gas-liquid chromatographic (GLC) instrumental and chromatographic system for the separation and flame ionization detection of the amino acid derivatives. In studies on the recovery of amino acids added to lunar fines, low recoveries were obtained when 10 ng of each amino acid were added to 50 mg of virgin fines, but the subsequent addition of 50 ng of each to the previously extracted sample resulted in much higher recoveries.
Detection limits were between 300 pg and 1 ng for different amino acids, in an analysis by gas-liquid chromatography of water extracts from Apollo 14 lunar fines in which amino acids were converted to their N-trifluoro-acetyl-n-butyl esters. Initial analyses of water and HCl extracts of sample 14240 and 14298 samples showed no amino acids above background levels.
An analytical model for separated airfoil flows is presented which is based on experimentally observed physical phenomena. These include a free stagnation point aft of the airfoil and a standing vortex in the separated region. A computer program is described which iteratively matches the outer potential flow, the airfoil turbulent boundary layer, the separated jet entrainment, mass conservation in the separated bubble, and the rear stagnation pressure. Separation location and pressure are not specified a priori. Results are presented for surface pressure coefficient and compared with experiment for three angles of attack for a GA(W)-1, 17% thick airfoil.
Activities reported include: (1) the construction of a thermocouple with a protective coating or sheath for obtaining reliable temperature measurements of gases in the base region; (2) the construction of a multiposition back step plate for trough tests for step sizes 0.63 cm and 0.25 cm; and (3) flow pattern studies with the 6.35 cm trough to determine the complex flow resulting when a shear layer reattaches in a corner. Reliable ignition and stable burning were achieved at Mach = 2 for the 0.63 cm trough. Photographs of side wall oil streaks for M=2 flow, and some static and pitot pressure results at M = 3 are included.
In studies of multijet rockets, failure to find solutions when using a reasonable mathematical model led to performing airflow tests of supersonic circular jets exhausting into noncircular channels of various sizes. Amomalous base pressure trends and puzzling wall pressure patterns resulted. A flow-into-the-corner test was then devised, using a 90 deg trough behind a back step. A vortex pair, generated by the unsymmetrical squeezing of the shear layer, was found to fill much of the corner and provided an explanation for the peculiar results of the previous tests. No previous reference could be found to this phenomenon. Further experiments of bodies with base fins defined the conditions for creating this trough vortex effect. The trough vortex was used successfully as an igniter/flameholder for combustion of hydrogen in a supersonic stream. Other applications are suggested for base drag reduction, base burning and for providing guidance to avoid errors in half-plane wind tunnel model testing.
The trough vortex phenomena is used for combustion of hydrogen in a supersonic air stream. This was done in small sizes suitable for igniters in supersonic combustion ramjets so long as the boundary layer displacement thickness is less than 25% of the trough step height. A simple electric spark, properly positioned, ignites the hydrogen in the trough corner. The resulting flame is self sustaining and reignitable. Hydrogen can be injected at the base wall or immediately upstream of the trough. The hydrogen is introduced at low velocity to permit it to be drawn into the corner vortex system and thus experience a long residence time in the combustion region. The igniters can be placed on a skewed back step for angles at least up to 30 deg. without affecting the igniter performance significantly. Certain metals (platinum, copper) act catalytically to improve ignition.
Computer modelling was used with two different designs of an advanced turbofan-ramjet in order to derive performance predictions. The engine would enable an aircraft to take-off, accelerate to Mach 5.0, and climb to 90,000 ft. The two concepts included a turbofan with a ramjet annularly wrapped around it and a side-by-side configuration with the ramjet having a rectangular shape and mounted alongside the turbofan. The studies were performed to model weight, length, fuel efficiency, and the requirements of the thrust/drag ratio to exceed unity over the entire flight path. LH2 would be used for fuel and to regeneratively cool the combustion chamber. Turbofan operation with and without afterburner and with and without the ramjet inlet open were examined, as were variable areas for the burners. A side-by-side configuration displayed the best performance predictions, with a ramjet mass flow being 75 percent that of the turbofan and maximum temperatures being equal.
This paper contains a discussion of the comparison of analytical and experimental results of the dynamic response of a flat rectangular plate subjected to electro-impulse type deicing forces. Early attempts in this correlation have been hampered by the complex leading edge geometries of the airfoils tested to date. The lack of a suitable analytical model for a typical leading edge structure has prompted these preliminary investigations of simple geometries, beginning with a flat plate and to be followed by a semi-cylindrical leading edge representation. The comparisons reported herein are thus limited to rectangular plate models. Project plans and icing tunnel results are given.
A joint University-Industry project has been sponsored by NASA Lewis Research Center to develop the Electro-Impulse method for de-icing aircraft. The program has consisted of basic analyses, laboratory testing, icing tunnel tests, and flight tests. During the past two years, the EIDI system has been tested and refined, and has been shown to be a low-energy, highly reliable de-icing system for a wide range of conditions. This paper gives a brief review of conditions. This paper gives a brief review of the basic principles, the development history, and results of recent flight tests by NASA and by Cessna Aircraft Company.
A brief review is given of four earlier tests in the NASA Lewis Icing Research Tunnel and of flight tests in NASA's Icing Research Aircraft and in a Cessna 206 airplane. Details are given of recent icing tunnel tests of thicker-skinned wings, a Gates Learjet, a composite leading edge, and a Boeing 767, and of a Falcon Fanjet engine inlet. These were tested at speeds from 87 to 220 knots, air temperatures from -2 to -15 C, LWC values of 0.6 to 2.4 grams/cu meter, and median droplet diameters from 12 to 20 microns. Energy requirements are reported, as well as conclusions from comparisons of several Electro-Impulse De-Icing coil system designs. Fundamental studies of the structural dynamics and ice shedding of a 12.7 cm (5 inch) diameter semicylinder are described. Some potential problem areas are discussed: fatigue of skin and coil mountings, system weight and cost, electro-magnetic interference and noise.
The application of electromagnetic impulse deicing (EIDI) systems to turbofan engine inlets on business aircraft has been investigated experimentally. The tests were performed in the Icing Research Tunnel at NASA's Lewis Research Center. The deicing system testbed was a Falcon Fanjet 20 engine nacelle. The effectiveness of various deicing coil configurations and mount designs were compared, and design parameters were developed specifically for EIDI systems in turbofan engines. Flight tests were also carried out at altitudes in the range 3000-6000 ft corresponding to a temperature range of -3 to -8 C. It is shown that the ice particles removed from the engine inlet by the deicing system were small enough for the engine to ingest. Tentative design specifications are given with respect to the optimum coil configuration, and operating power of a EIDI production candidate.
De-icing of aircraft by using the electro-magnetic impulse phenomenon was proposed and demonstrated in several European countries. However, it is not available as a developed system due to lack of research on the basic physical mechanisms and necessary design parameters. The de-icing is accomplished by rapidly discharging high voltage capacitors into a wire coil rigidly supported just inside the aircraft skin. Induced eddy currents in the skin create a repulsive force resulting in a hammer-like force which cracks, de-bonds, and expels ice on the skin surface. The promised advantages are very low energy, high reliability of de-icing, and low maintenance. Three years of Electo-Impulse De-icing (EIDI) research is summarized and the analytical studies and results of testing done in the laboratory, in the NASA Icing Research Tunnel, and in flight are presented. If properly designed, EIDI was demonstrated to be an effective and practical ice protection system for small aircraft, turbojet engine inlets, elements of transport aircraft, and shows promise for use on helicopter rotor blades. Included are practical techniques of fabrication of impulse coils and their mountings. The use of EIDI with nonmetallic surface materials is also described.
The Altitude Wind Tunnel at the NASA-Lewis facility is being proposed for a refurbishment and moderization. Two major changes are: (1) the increasing of the test section Mach number to 0.90, and (2) the addition of spray nozzles to provide simulation of flight in icing clouds. Features to be retained are the simulation of atmospheric temperature and pressure to 50,000 foot altitude and provision for full-scale aircraft engine operation by the exhausting of the aircraft combustion gases and ingestion of air to replace that used in combustion. The first change required a re-design of the turning vanes in the two corners downstream of the test section due to the higher Mach number at the corners. The second change threatens the operation of the turning vanes by the expected ice build-up, particulary on the first-corner vanes. De-icing by heat has two drawbacks: (1) an extremely large amount of heat is required, and (2) the melted ice would tend to collect as ice on some other surfaces in the tunnel, namely, the tunnel propellers and the cooling coils. An alternate de-icing method had been under development for three years under NASA-Lewis grants to the Wichita State University. This report describes the electro-impulse de-icing (EIDI) method and the testing work done to assess its applicability to wind tunnel turning vane de-icing. Tests were conducted in the structural dynamics laboratory and in the NASA Icing Research Tunnel. Good ice protection was achieved at lower power consumption and at a wide range of tunnel operations conditions. Recommendations for design and construction of the system for this application of the EIDI method are given.
Basic principles and parameters for a system to deice aircraft with electromagnetic impulses are described. The physical basis for deicing by such impulses is explained, and the requirements involved in the electrodynamic design, structural dynamic design, and system design are discussed. Some manufacturing and testing problems and techniques are described.
An efficient and accurate method is described for extracting water droplet impingement efficiency data from dye impregnated blotter paper samples obtained by employing a dye-tracer technique in wind tunnel spray tests. The method is based on laser reflectance spectroscopy. A brief description of the test method, instrumentation, and data reduction system is also presented. Preliminary test results and analyses are included for a cylinder and a 65(2)015 airfoil.