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  • Polymer and Materials Science  (10,782)
  • General Chemistry  (7,564)
  • PROPULSION SYSTEMS  (1,456)
  • Cell & Developmental Biology
  • 1970-1974  (17,938)
  • 1910-1914  (2,932)
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  • 1
    Publication Date: 2006-01-11
    Description: An important noise source in a drilling plant is Diesel engine exhaust. In order to reduce this noise, a reactive silencer of the derivative resonator type was proposed, calculated from the acoustic and design point of view and applied. As a result of applying such a silencer on the exhaust conduit of a Diesel engine the noise level dropped down to 18 db.
    Keywords: PROPULSION SYSTEMS
    Type: The 4th Natl. Conf. on Acoustics, Vol. 1A (NASA-TT-F-15375); p 156-162
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  • 2
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    In:  Other Sources
    Publication Date: 2011-08-16
    Description: The state of the art of electric propulsion technology is discussed with a review of the many difficult engineering problems which must be resolved before the incorporation of electric propulsion in space missions. Interferences with the spacecraft and its scientific instrumentation, conducted and radiated electromagnetic interferences during the switching and processing of large electrical loads, and mercury and other eroded materials deposition on the spacecraft are indicated among the engineering and design challenges to be overcome and resolved.
    Keywords: PROPULSION SYSTEMS
    Type: AIAA Student Journal; 12; Apr. 197
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  • 3
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    In:  Other Sources
    Publication Date: 2011-08-16
    Description: Engine designs resulting from the NASA sponsored quiet engine program are described in terms of their possible use on the DC-8. These engines have high bypass ratios for low jet noise and low-speed, low-noise fans. Experiments were carried out on various fan designs with different tip speeds. The advantage of low tip speeds and high lift coefficients has been demonstrated. The engines are equipped with flight-standard nacelles with acoustic suppression. It is estimated that these improvements will reduce DC-8 noise by about 26EPNdB-29EPNdB.
    Keywords: PROPULSION SYSTEMS
    Type: Flight International; 106; July 25
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  • 4
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    In:  Other Sources
    Publication Date: 2011-08-16
    Description: An assessment of the present state of the art in noise reduction technology indicates that this technology has the potential for effectively attaining this goal - a conclusion that is in apparent conflict with the frequently voiced complaints on intolerable noise levels near airports. Measures are suggested for a more vigorous implementation of available technology in practice to combat the aircraft noise problem.
    Keywords: PROPULSION SYSTEMS
    Type: Astronautics and Aeronautics; 12; Jan. 197
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  • 5
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    In:  Other Sources
    Publication Date: 2011-08-16
    Description: Procedures for protecting aircraft engines from impact damage due to sand and dust particles are discussed. One system is to develop a filtration system to remove most of the solid particles. Another method is to design blade materials and blade protection coatings. The concepts of erosion within aircraft engines are explained. Curves are presented to show the effect of impingement angle on ductile and elastic modes of erosion. Engine sand and dust ingestion tests were conducted on T63-A-5A engines and the results are produced in graph form. The results of tests conducted in a dynamic erosion facility are analyzed.
    Keywords: PROPULSION SYSTEMS
    Type: NBS The Role of Cavitation in Mech. Failures; p 129-145
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  • 6
    Publication Date: 2012-05-22
    Description: The severity of proposed noise goals for STOL systems has resulted in a new design approach for aircraft propulsion systems. It has become necessary to consider the influence of the noise goal on the design of engine components, engine systems, and the integrated nacelle, separately and collectively, from the onset of the design effort. This integrated system design approach is required in order to effect an optimization of the propulsion and aircraft system. Results from extensive design studies and pertinent test programs are presented which show the effect of noise specifications on component and system design, and the trade-offs possible of noise versus configuration and performance. The design optimization process of propulsion systems for powered lift systems is presented beginning with the component level and proceeding through to the final integrated propulsion system. Designs are presented which are capable of meeting future STOL noise regulations and the performance, installation and economic penalties are assessed as a function of noise level.
    Keywords: PROPULSION SYSTEMS
    Type: AGARD V(STOL Propulsion Systems; 17 p
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  • 7
    Publication Date: 2019-06-27
    Description: Because of the relatively high takeoff speeds of supersonic transport aircraft, it is important to know whether the flight velocity effects the noise level of suppressor nozzles. To investigate this, a modified F-106B aircraft was used to conduct a series of flyover and static tests on a 12-chute suppressor installed on an uncooled plug nozzle. Comparison of flyover and static spectra indicated that flight velocity adversely affected noise suppressions of the 12-chute configurations.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-2918 , E-7449
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  • 8
    Publication Date: 2019-06-27
    Description: A computer program is presented and described that generates jet engine inlet flow contour maps and inlet flow distortion parameters. The program input consists of an array of measurements describing the flow conditions at the engine inlet. User-defined distortion parameters may be calculated.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-2967 , E-7572
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  • 9
    Publication Date: 2019-06-27
    Description: Plug nozzles with two types of 40-spoke noise suppressor were tested at free-stream Mach numbers from 0 to 0.45 and over a range of nozzle pressure ratios from 1.5 to 4.0. In additon, an unsuppressed plug nozzle and a Supersonic Tunnel Association nozzle were also tested to provide baseline levels of thrust performance. The unsuppressed plug nozzle had an efficiency of 98 percent at an assumed takeoff pressure ratio of 3.0 and at Mach 0.36. At the same condition the suppressor nozzles had efficiencies of approximately 83.5 percent.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-2951 , E-7541
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  • 10
    Publication Date: 2019-06-27
    Description: Because of the relatively high takeoff speeds of supersonic transport aircraft, it is important to know if the flight velocity affects the noise level of suppressor nozzles. To investigate this, a modified F-106B aircraft was used to conduct a series of flyover and static tests on a 48-tube suppressor installed on an uncooled plug nozzle. Comparison of flyover and static spectra indicated that flight velocity had little effect on the noise suppression of the 48-tube suppressor configuration. However, flight velocity adversely affected noise suppression of the 48-tube suppressor with an acoustic shroud and plug installed.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-2919 , E-7513
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  • 11
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    In:  CASI
    Publication Date: 2019-06-27
    Description: The SERT C (Space Electric Rocket Test - C) project study defines a spacecraft mission that would demonstrate the technology readiness of ion thruster systems for primary propulsion and station keeping applications. As a low cost precursor, SERT C develops the components and systems required for subsequent Solar Electric Propulsion (SEP) applications. The SERT C mission requirements and preliminary spacecraft and subsystem design are described.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-71508
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  • 12
    Publication Date: 2019-06-27
    Description: The effect of film coolant ejection from the pressure side of a stator blade was determined in a two-dimensional cascade. Stator exit surveys were made for each of six rows of coolant holes. Successive multirow tests were made with two, three, four, five, and six rows of coolant holes open. The results of the multirow tests are compared with the predicted multirow performance obtained by adding the single-row data. Results are presented in terms of stator primary-air efficiency as a function of coolant fraction.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-3045 , E-7817
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  • 13
    Publication Date: 2019-06-27
    Description: An experimental investigation was conducted to determine the thermal effectiveness of an aluminized Mylar-silk net insulation system containing up to 160 layers. The experimentally measured heat flux was compared with results predicted by using (1) a previously developed semi-empirical equation and (2) an effective-thermal-conductivity value. All tests were conducted at a nominal hot-boundary temperature of 294 K (530 R) with liquid hydrogen as the heat sink. The experimental results show that the insulation performed as expected and that both the semi-empirical equation and effective thermal conductivity of a small number of layers were adequate in predicting the thermal performance of a large number of layers of insulation.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-7659 , E-7815
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  • 14
    Publication Date: 2019-06-27
    Description: The work is described which was performed on the design, analysis and performance of a 20 kW rollup solar array capable of meeting the design requirements of a solar electric spacecraft for the 1980 Encke rendezvous mission. To meet the high power requirements of the proposed electric propulsion mission, solar arrays on the order of 186.6 sq m were defined. Because of the large weights involved with arrays of this size, consideration of array configurations is limited to lightweight, large area concepts with maximum power-to-weight ratios. Items covered include solar array requirements and constraints, array concept selection and rationale, structural and electrical design considerations, and reliability considerations.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-136860 , JPL-TM-33-668
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  • 15
    Publication Date: 2019-06-27
    Description: Tests were conducted to determine the performance of a hydrogen burner used to produce a test gas that simulates air entering a scramjet combustor at various flight conditions. The test gas simulates air in that it duplicates the total temperature, total pressure, and the volume fraction of oxygen of air at flight conditions. The main objective of the tests was to determine the performance of the burner as a function of the effective exhaust port area. The conclusions were: (1) pressure oscillations of the chugging type were reduced in amplitude to plus or minus 2 percent of the mean pressure level by proper sizing of hydrogen, oxygen, and air injector flow areas; (2) combustion efficiency remained essentially constant as the exhaust port area was increased by a factor of 3.4; (3) the mean total temperature determined from integrating the exit radial gas property profiles was within plus or minus 5 percent of the theoretical bulk total temperature; (4) the measured exit total temperature profile had a local peak temperature more than 30 percent greater than the theoretical bulk total temperature; and (5) measured heat transfer to the burner liner was 75 percent of that predicted by theory based on a flat radial temperature profile.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-7567 , L-9363
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  • 16
    Publication Date: 2019-06-27
    Description: The experimental test program results of a 4 1/2-stage turbine with a very high stage loading factor are presented. A four-stage turbine was tested with and without outlet turning vanes. The 4 1/2-stage turbine achieved a design point total-to-total efficiency of 0.853. The outlet turning vane design point performance was 0.4 percent of the overall 4 1/2-stage turbine efficiency. Tests were conducted at various levels of Reynolds number and indicated decreases in turbine efficiency and equivalent weight flow with decreasing Reynolds number.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-2363 , R73AEG163
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  • 17
    Publication Date: 2019-06-27
    Description: An axial flow compressor stage, having tandem airfoil blading, was designed for zero rotor prewhirl, constant rotor work across the span, and axial discharge flow. The stage was designed to produce a pressure ratio of 1.265 at a rotor tip velocity of 757 ft/sec. The rotor has an inlet hub/tip ratio of 0.8. The design procedure accounted for the rotor inlet boundary layer and included the effects of axial velocity ratio and secondary flow on blade row performance. The objectives of this experimental program were (1) to obtain performance with uniform and distorted inlet flow for comparison with the performance of a stage consisting of single-airfoil blading designed for the same vector diagrams and (2) to evaluate the effectiveness of accounting for the inlet boundary layer, axial velocity ratio, and secondary flows in the stage design.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134529 , FR-5955
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  • 18
    Publication Date: 2019-06-27
    Description: Results of supersonic mixing and combustion tests performed with two simple strut injector configurations, one with parallel injectors and one with perpendicular injectors, are presented and analyzed. Good agreement is obtained between static pressure measured on the duct wall downstream of the strut injectors and distributions obtained from one-dimensional calculations. Measured duct heat load agrees with results of the one-dimensional calculations for moderate amounts of reaction, but is underestimated when large separated regions occur near the injection location. For the parallel injection strut, good agreement is obtained between the shape of the injected fuel distribution inferred from gas sample measurements at the duct exit and the distribution calculated with a multiple-jet mixing theory. The overall fraction of injected fuel reacted in the multiple-jet calculation closely matches the amount of fuel reaction necessary to match static pressure with the one-dimensional calculation. Gas sample measurements with the perpendicular injection strut also give results consistent with the amount of fuel reaction in the one-dimensional calculation.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-7581 , L-9428
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  • 19
    Publication Date: 2019-06-27
    Keywords: PROPULSION SYSTEMS
    Type: Acoustical Society of America; vol. 56
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  • 20
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    In:  CASI
    Publication Date: 2019-06-27
    Description: A structural integrated ion thruster with 8-cm beam diameter (SIT-8) was developed for attitude control and stationkeeping of synchronous satellites. As optimized, the system demonstrates a thrust T=1.14 mlb (not corrected for beam V sub B = 1200 V (I sub sp = 2200 sec) total propellant utilization efficiency nu sub u = 59.8% (is approximately 72% without auxiliary pulse-igniter electrode), and electrical efficiency n sub E 61.9%. The thruster incorporates a wire-mesh anode and tantalum cover surfaces to control discharge chamber flake formation and employs an auxiliary pulse-igniter electrode for hollow-cathode ignition. When the SIT-8 is integrated with the compatible SIT-5 propellant tankage, the system envelope is 35 cm long by 13 cm flange bolt circle with a mass of 9.8 kg including 6.8 kg of mercury propellant. Two thrust vectoring systems which generate beam deflections in two orthogonal directions were also developed under the program and tested with the 8-cm thruster. One system vectors the beam over + or - 10 degrees by gimbaling of the entire thruster (not including tankage), while the other system vectors the beam over + or - 7 degrees by translating the accel electrode relative to the screen electrode.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134685
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  • 21
    Publication Date: 2019-06-27
    Description: A 51-cm-diam. model of a short-haul fan stage was tested. Surveys of the air flow conditions ahead of the rotor, between the rotor and stator, and behind the stator were made over the stable operating range of the stage. At the design speed of 213.3 m/sec and weight flow of 26.7 kg/sec, the stage pressure ratio and efficiency are 1.18 and 0.87, respectively. The rotor peak efficiency of 0.92 occurred at flow rate of 30.5 kg/sec. Peak stage efficiency of 0.09 was obtained at 110 percent speed at a pressure ratio of 1.218 and a weight flow of 30.2 kg/sec. Maximum stage pressure ratio is 1.269 at 120 percent speed.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-3101 , E-7791
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  • 22
    Publication Date: 2019-06-27
    Description: The characteristics and performance of the orbit maneuvering engine for the space shuttle are discussed. Emphasis is placed on the regeneratively cooled thrust chamber of the engine. Tests were conducted to determine engine operating parameters during the start, shutdown, and restart. Characteristics of the integrated thrust chamber and the performance and thermal conditions for blowdown operation without supplementary boundary layer cooling were investigated. The results of the test program are presented.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-140250 , ASR74-229
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  • 23
    Publication Date: 2019-06-27
    Description: A solid propellant composition of improved efficiency is described which includes an oxidizer containing ammonium perchlorate, and a powered metal fuel, preferably aluminum or beryllium, in the form of a composite. The metal fuel is contained in the crystalline lattice framework of the oxidizer, as well as within the oxidizer particles, and is disposed in the interstices between the oxidizer particles of the composition. The propellant composition is produced by a process comprising the crystallization of ammonium perchlorate in water, in the presence of finely divided aluminum or beryllium. A suitable binder is incorporated in the propellant composition to bind the individual particles of metal with the particles of oxidizer containing occluded metal.
    Keywords: PROPULSION SYSTEMS
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  • 24
    Publication Date: 2019-06-27
    Description: The successful hot fire demonstration of a pulsing liquid hydrogen/liquid oxygen and gaseous hydrogen/liquid oxygen attitude control propulsion system thruster is described. The test was the result of research to develop a simple, lightweight, and high performance reaction control system without the traditional requirements for extensive periods of engine thermal conditioning, or the use of complex equipment to convert both liquid propellants to gas prior to delivery to the engine. Significant departures from conventional injector design practice were employed to achieve an operable design. The work discussed includes thermal and injector manifold priming analyses, subscale injector chilldown tests, and 168 full scale and 550 N (1250 lbF) rocket engine tests. Ignition experiments, at propellant temperatures ranging from cryogenic to ambient, led to the generation of a universal spark ignition system which can reliably ignite an engine when supplied with liquid, two phase, or gaseous propellants. Electrical power requirements for spark igniter are very low.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134655
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  • 25
    Publication Date: 2019-06-27
    Description: The performance and heat transfer characteristics of a doublet element type injector for the space shuttle orbiter maneuvering engine thrust chamber were investigated. Ths stability characteristics were evaluated over a range of chamber pressures and mixture ratios. The specific objectives of the test were: (1) to determine whether stability has been influenced by injection of boundary layer coolant across the cavity entrance, (2) if the injector is stable, to determine the minimum cavity area required to maintain stability, and (3) if the injector is unstable, to determine the effects of entrance geometry and increased area on stability.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134359 , ASR74-204
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  • 26
    Publication Date: 2019-06-27
    Description: This study had the objectives of determining the most promising conventional and variable cycle engine types; the effect of design cruise Mach number (2.2, 2.7 and 3.2) on a commercial supersonic transport; effect of advanced engine technology on the choice of engine cycle; and effect of utilizing hydrogen as the engine fuel. The technology required for the engines was defined, and the levels of development to ensure availability of this technology in advanced aircraft propulsion systems were assessed. No clearcut best conventional or variable cycle engine was identified. The dry bypass turbojet and the duct burning turbofans were initially selected as the best conventional engines, but later results, utilizing augmentation at takeoff, added the mixed-flow augmented turbofan as a promising contender. The modulating air flow, three-rotor variable cycle engine identified the performance features desired from VCE concepts (elimination of inlet drag and reduction in afterbody drag), but was a very heavy and complex engine.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-143634 , R74AEG330
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  • 27
    Publication Date: 2019-06-27
    Description: Results of an experimental program conducted on a mixed-compression inlet coupled to a turbofan engine are presented. Open-loop frequency response data are presented that show the response of shock position (as measured by an average inlet static pressure) to sinusoidal airflow disturbances produced at the compressor face station. Also presented are results showing the effect of different passive terminations (a choke plate or a long duct) on the characteristics of the inlet. Transfer functions obtained by using experimental data are presented and compared to the experimental data. Closed-loop frequency response of shock position (with a proportional-plus-integral controller) is presented. In addition, transient data are presented that show the unstart-restart characteristics of the inlet.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-3104 , E-7924
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  • 28
    Publication Date: 2019-06-27
    Description: Computer programs for analyzing the acoustic properties of turbomachinery with ducted flow were developed. The models include component interactions and rotor alone. Test case results determined from the computer programs are presented.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-132445 , D6-43296-3
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  • 29
    Publication Date: 2019-06-27
    Description: The experimental results are summarized that were obtained in the development of a Teflon solid propellant pulsed plasma thruster. The feasibility was established of storing and feeding solid propellant in the form of an open circular loop into an operational thruster. This technique was verified to be practical by feeding over 20 inches of Teflon into a micro-thruster over an accumulated life test of 1858 hours. High energy density capacitors were evaluated under vacuum conditions when the capacitor is coupled directly to a plasma thruster. Numerous early capacitor failures were encountered. It was concluded that essentially all of the failures encountered in a vacuum environment are due to an internal electrical breakdown that will occur inside a capacitor that is not truly hermetically sealed. A steady input power significantly in excess of 130 watts can safely be tolerated if heat conduction can be provided to a sink whose temperature is about 16 C. A vacuum life test of the capacitor bank was carried out while discharging into a milli-lb. (milli-Newton) type pulsed plasma thruster. More than 1500 hours of vacuum testing of this milli-Newton type system has been accumulated without any capacitor problems. Recommendations are made for future capacitor designs.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-139272 , MS172R0001
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  • 30
    Publication Date: 2019-06-27
    Description: Analytical models were developed for computing the periodic sound pressures of subsonic fans in an infinite hardwall annular duct with uniform flow. The computer programs are described which are used for numerical computations of sound pressure mode amplitudes. The data are applied to the acoustic properties of turbomachinery.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-132444 , D6-43296-2
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  • 31
    Publication Date: 2019-06-27
    Description: The analytic models developed for computing the periodic sound pressure of subsonic fans and compressors in an infinite, hardwall annular duct with uniform flow are described. The basic sound-generating mechanism is the scattering into sound waves of velocity disturbances appearing to the rotor or stator blades as a series of harmonic gusts. The models include component interactions and rotor alone.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-132443 , D6-43296-1
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  • 32
    Publication Date: 2019-06-27
    Description: A significant effort within the NASA Quiet Engine Program has been devoted to acoustical evaluation at the Lewis Research Center noise test facility of a family of full-scale fans. This report, documents the noise results obtained with fan A - a 1.5-pressure-ratio, 1160-ft/sec-tip-speed fan. The fan is described and some aerodynamic operating data are given. Far-field noise around the fan was measured for a variety of configurations pertaining to acoustical treatment and over a range of operating conditions. Complete results of 1/3-octave band analysis of the data are presented in tabular form. Included also are power spectra and sideline perceived noise levels. Some representative 1/3-octave band data are presented graphically, and sample graphs of continuous narrow-band spectra are also provided.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-3066 , E-7857
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  • 33
    Publication Date: 2019-06-27
    Description: An experimental wind tunnel investigation was conducted to determine the effects of time-variant distortions produced in a supersonic inlet on a J85-GE-13 turbojet engine. Results are presented principally in terms of instantaneous distortion amplitudes and total-pressure contours measured through compressor stall. They indicate that, although a time-averaged distortion may be far from a stall-inducing value, corresponding instantaneous distortion amplitudes can approach or exceed this value. Seven engine stall events were studied. In six of these events instantaneous distortion peaks of sufficient magnitude to cause stall were observed in the time period shortly before stall was detected.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-3002 , E-6294
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  • 34
    Publication Date: 2019-06-27
    Description: The particle motion in two-dimensional free and forced inward flowing vortices is considered. A particle in such a flow field experiences a balance between the aerodynamic drag forces that tend to drive erosive particles toward the axis, and centrifugal forces that prevent these particles from traveling toward the axis. Results predict that certain sizes of particles will achieve a stable orbit about the turbine axis in the inward flowing free vortex. In this condition, the radial drag force is equal to the centrifugal force. The sizes of particles that will achieve a stable orbit is shown to be related to the gas flow velocity diagram at a particular radius. A second analysis yields a description of particle sizes that will experience a centrifugal force that is greater than the radial component of the aerodynamic drag force for a more general type of particle motion.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134616
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  • 35
    Publication Date: 2019-06-27
    Description: A study program for conceptual design studies of remote lift and lift/cruise fan systems to meet the requirements of military V/STOL aircraft was conducted. Parametric performance and design data are presented for fans covering a range of pressure ratios, including both single and two stage fan concepts. The gas generator selected for these fan systems was the J101-GE-100 engine. Noise generation and transient response were determined for selected fan systems.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134636 , R74AEG283
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  • 36
    Publication Date: 2019-06-27
    Description: The use of heat conduction flame generated in a premixed supersonic stream is discussed. It is shown that the flame is controlled initially by heat conduction and then by chemical reaction. Such a flame is shorter than the diffusion type of flame and therefore it requires a much shorter burner. The mixing is obtained by injecting the hydrogen in the inlet. Then the inlet can be cooled by film cooling.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-132446
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  • 37
    Publication Date: 2019-06-27
    Description: An analysis to predict the pressurant gas requirements for the discharge of cryogenic liquid propellants from storage tanks is presented, along with an algorithm and two computer programs. One program deals with the pressurization (ramp) phase of bringing the propellant tank up to its operating pressure. The method of analysis involves a numerical solution of the temperature and velocity functions for the tank ullage at a discrete set of points in time and space. The input requirements of the program are the initial ullage conditions, the initial temperature and pressure of the pressurant gas, and the time for the expulsion or the ramp. Computations are performed which determine the heat transfer between the ullage gas and the tank wall. Heat transfer to the liquid interface and to the hardware components may be included in the analysis. The program output includes predictions of mass of pressurant required, total energy transfer, and wall and ullage temperatures. The analysis, the algorithm, a complete description of input and output, and the FORTRAN 4 program listings are presented. Sample cases are included to illustrate use of the programs.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-7504 , E-7654
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  • 38
    Publication Date: 2019-06-27
    Description: The combustion characteristics of hypergolic propellants are described. A research project was conducted to determine if the reaction control system engine propellants on Apollo spacecraft undergo explosive reaction when subjected to conditions present in the engine at the time of ignition. Combustion characteristics pertinent to the hard-start phenomenon are considered. The thermal stability of frozen mixtures of hydrazine-based fuels with nitrogen tetroxide was analyzed. Results of the tests are presented in the form of tables and graphs.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134314 , BM-IR-1646-VOL-2
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  • 39
    Publication Date: 2019-06-27
    Description: The results of an airline study of the economic effects of propulsion system technology on current and future transport aircraft are presented. This report represents the results of a detailed study of propulsion system operating economics. The study has four major parts: (1) a detailed analysis of current propulsion system maintenance with respect to the material and labor costs encountered versus years in service and the design characteristics of the major elements of the propulsion system of the B707, b727, and B747. (2) an analysis of the economic impact of a future representative 1979 propulsion system is presented with emphasis on depreciation of investment, fuel costs and maintenance costs developed on the basis of the analysis of the historical trends observed. (3) recommendations concerning improved methods of forecasting the maintenance cost of future propulsion systems are presented. A detailed method based on the summation of the projected labor and material repair costs for each major engine module and its installation along with a shorter form suitable for quick, less detailed analysis are presented, and (4) recommendations concerning areas where additional technology is needed to improve the economics of future commercial propulsion systems are presented along with the suggested economic benefits available from such advanced technology efforts.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134645
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  • 40
    Publication Date: 2019-06-27
    Description: An analytic model is developed to predict pressure and flow transients in a gaseous hydrogen-oxygen reaction control rocket engine feed system. The one-dimensional equations of momentum and continuity are reduced by the method of characteristics from partial derivatives to a set of total derivatives which describe the state properties along the feedline. System components, e.g., valves, manifolds, and injectors are represented by pseudo steady-state relations at discrete junctions in the system. Solutions were effected by a FORTRAN IV program on an IBM 360/65. The results indicate the relative effect of manifold volume, combustion lag time, feedline pressure fluctuations, propellant temperature, and feedline length on the chamber pressure transient. The analytical combustion model is verified by good correlation between predicted and observed chamber pressure transients. The developed model enables a rocket designer to vary the design parameters analytically to obtain stable combustion for a particular mode of operation which is prescribed by mission objectives.
    Keywords: PROPULSION SYSTEMS
    Type: Computers and Fluids; 2; Aug. 197
    Format: text
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  • 41
    facet.materialart.
    Unknown
    In:  Other Sources
    Publication Date: 2019-06-27
    Description: The feasibility of electrical power generation with an ambient temperature liquid-metal MHD separator cycle is demonstrated by tests in which a NaK-nitrogen LM-MHD converter was operated at nozzle inlet pressures ranging from 100 to 165 N/sq cm, NaK flow rates from 46 to 72 kg/sec, and nitrogen flow rates from 2.4 to 3.8 kg/sec. The generator was operated as an eight-phase linear induction generator, with two of the eight phases providing magnetic field compensation to minimized electrical end losses at the generator channel inlet and exit.
    Keywords: PROPULSION SYSTEMS
    Type: AIAA Journal; 12; Jan. 197
    Format: text
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  • 42
    Publication Date: 2019-06-27
    Description: For abstract, see N74-35204.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-140309 , R-9557-1
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  • 43
    Publication Date: 2019-06-27
    Description: An experimental test program was performed to demonstrate the durability of a beryllium INTEREGEN rocket engine when operating under conditions simulating the space shuttle reaction control system. A vibration simulator was exposed to the equivalent of 100 missions of X, Y, and Z axes random vibration to demonstrate the integrity of the recently developed injector-to-chamber braze joint. An off-limits engine was hot fired under extreme conditions of mixture ratio, chamber pressure, and orifice plugging. A durability engine was exposed to six environmental cycles interspersed with hot-fire tests without intermediate cleaning, service, or maintenance. Results from this program indicate the ability of the beryllium INTEREGEN engine concept to meet the operational requirements of the space shuttle reaction control system.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-140308 , R-9557
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  • 44
    Publication Date: 2019-06-27
    Description: A two-dimensional inlet and a turbojet engine were investigated in a Mach 0.4 stream so as to ingest the tip vortex of a forward mounted wing. Results show that ingestion of a wing tip vortex by a turbojet engine can cause a large reduction in engine stall margin. The loss in stall compressor pressure ratio was primarily dependent on vortex location and rotational direction and not on total-pressure variations across the compressor face.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-71610 , E-8091
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  • 45
    Publication Date: 2019-06-27
    Description: Experimental data were obtained with a 21.59 cm (8.5 in.) diameter cold-flow model in a static altitude facility to determine the thrust and pumping characteristics of several variations of a ram-air-cooled plug nozzle. Tests were conducted over a range of nozzle pressure ratios simulating supersonic cruise and takeoff conditions. Primary throat area was also varied to simulate afterburner on and off. Effect of plug size, outer shroud length, primary nozzle geometry, and varying amounts of secondary flow were investigated. At a supersonic cruise pressure ratio of 27, nozzle efficiencies were 99.7 percent for the best configurations.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-3110 , E-7975
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  • 46
    Publication Date: 2019-06-27
    Description: An evaluation of the technology and the development of a preliminary design for a 25 kilowatt solar array system for solar electric propulsion are discussed. The solar array has a power to weight ratio of 65 watts per kilogram. The solar array system is composed of two wings. Each wing consists of a solar array blanket, a blanket launch storage container, an extension/retraction mast assembly, a blanket tensioning system, an array electrical harness, and hardware for supporting the system for launch and in the operating position. The technology evaluation was performed to assess the applicable solar array state-of-the-art and to define supporting research necessary to achieve technology readiness for meeting the solar electric propulsion system solar array design requirements.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-120483 , LMSC-D384250
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  • 47
    facet.materialart.
    Unknown
    In:  CASI
    Publication Date: 2019-06-27
    Description: The flow field of a jet engine with an inlet vortex, which can pick up heavy debris from the ground and damage the engine, was simulated in a small water tunnel by means of the hydrogen bubble technique. It was found that the known engine inlet vortex is accompained by a vortex system, consisting of two inlet vortices (the ground based and the trailing one), secondary vortices, and ground vortices. Simulation of the ground effect by an inlet image proved that the inlet vortex feeds on free stream vorticity and can exist without the presence of a ground boundary layer. The structural form of the inlet vortex system was explained by a simple potential flow model, which showed the number, location, and the importance of the stagnation points. A retractable horizontal screen or an up-tilt of the engine is suggested as countermeasure against debris ingestion.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-140182
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  • 48
    facet.materialart.
    Unknown
    In:  CASI
    Publication Date: 2019-06-27
    Description: Criteria for the design and development of turbines for rocket engines to meet specific performance, and installation requirements are summarized. The total design problem, and design elements are identified, and the current technology pertaining to these elements is described. Recommended practices for achieving a successful design are included.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-SP-8110
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  • 49
    Publication Date: 2019-06-27
    Description: The operational flight instrumentation required for performance monitoring and fault detection are presented. Measurements by the burn through monitors are presented along with manifold and helium source pressures.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-140258 , LEC-3594 , JSC-09130
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  • 50
    Publication Date: 2019-06-27
    Description: A computer program, designated RETSCP, for the analysis of Rocket Engine Thermal Strain with Cyclic Plasticity is described. RETSCP is a finite element program which employs a three dimensional isoparametric element. The program treats elasto-plastic strain cycling including the effects of thermal and pressure loads and temperature dependent material properties. Theoretical aspects of the finite element method are discussed and the program logic is described. A RETSCP User's Manual is presented including sample case results.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134640
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  • 51
    Publication Date: 2019-06-27
    Description: A finite element stress analysis was performed for the film cooled throat section of an attitude control thruster. The anlaysis employed the RETSCP finite element computer program. The analysis included thermal and pressure loads, and the effects of temperature dependent material properties, to determine the strain range corresponding to the thruster operating cycle. The configuration and operating conditions considered, correspond to a flightweight integrated thruster assembly which was thrust pulse tested. The computed strain range was used in conjuction with Haynes 188 Universal Slopes minimum life data to predict throat section fatigue life. The computed number of cycles to failure was greater than the number of pulses to which the thruster was experimentally subjected without failure.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134641-VOL-2
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  • 52
    Publication Date: 2019-06-27
    Description: A three-dimensional finite element elasto-plastic strain analysis was performed for the throat section of a regeneratively cooled rocket combustion chamber. The analysis employed the RETSCP finite element computer program. The analysis included thermal and pressure loads, and the effects of temperature dependent material properties, to determine the strain range corresponding to the chamber operating cycle. The analysis was performed for chamber configuration and operating conditions corresponding to a hydrogen-oxygen combustion chamber which was fatigue tested to failure. The computed strain range at typical chamber operating conditions was used in conjunction with oxygen-free, high-conductivity (OHFC) copper isothermal fatigue test data to predict chamber low-cycle fatigue life.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134641-VOL-1
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  • 53
    Publication Date: 2019-06-27
    Description: A generalized analytical model and computer program have been developed to predict the frequency response of arbitrary liquid propellant feedline designs. The analytical model is based on an extension of an existing distributed parameter representation of a viscous fluid transmission line with laminar flow which was modified to include the effects of a turbulent mean flow. The effects of dissolved ullage gases, wall elasticity, localized gas or vapor bubbles, bellows, forced changes in length due to structural excitation, complex side branches, and structural mounting stiffness are also included. Each line component is written as a four-terminal, pressure-flow relationship in matrix form in the Laplace domain; the transfer function relating the pressure response at the line terminal (inducer inlet) to the external excitation is obtained in the computer program by sequential matrix substitution.
    Keywords: PROPULSION SYSTEMS
    Type: Journal of Spacecraft and Rockets; 11; Mar. 197
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  • 54
    Publication Date: 2019-06-27
    Description: The results are presented of a preliminary flight engine design study based on the Quiet Engine Program high-bypass, low-noise turbofan engines. Engine configurations, weight, noise characteristics, and performance over a range of flight conditions typical of a subsonic transport aircraft were considered. High and low tip speed engines in various acoustically treated nacelle configurations were included.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134660 , R74AEG327
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  • 55
    Publication Date: 2019-06-27
    Description: Four injector designs and two chamber profiles were experimentally evaluated for structural integrity, combustion efficiency, and resistance to combustion instabilities. Vacuum thrust measurements were used as a primary measure of combustion efficiency. Stability rating to test the sensitivity of the injectors to high frequency combustion was conducted, but not extensively. To map the boundary between stable operation and chugging instability, chamber pressure was throttled downward from 689.5 to 206.9 kN/sq m abs (100 to 30 psia). Best operational results were obtained with an injector configuration having no hydraulic swirlers, a 0.00102-m (0.040-in.) recessed FLOX tube, and a nonflared exit in the methane annulus. This injector design exhibited stable combustion and good integrity of hardware, and it exceeded the design goal efficiency (88 percent) at the 10 to 1 throttled condition.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-3094 , E-7911
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  • 56
    Publication Date: 2019-06-27
    Description: A computer program was developed to describe the performance of ramjet and scramjet cycles. The program performs one dimensional calculations of the equilibrium, real-gas internal flow properties of the engine. The program can be used for the following: (1) preliminary design calculation and (2) design analysis of internal flow properties corresponding to stipulated flow areas. Only the combustion of hydrogen in air is considered in this case.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-132454 , AP-1001-1
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  • 57
    Publication Date: 2019-06-27
    Description: An isolated 1.15 pressure ratio turbofan engine simulator was tested at Mach numbers from 0.6 to 0.85. At Mach 0.75 the net propulsive force of the fan and nacelle (excluding core thrust) was 73 percent of the ideal fan net thrust. Internal losses amounted to 7 percent, and external drag amounted to 20 percent of the ideal fan net thrust. External pressure and friction drag were about equal. The propulsive efficiency with a 90 percent efficient fan would have been 63 percent. For the aerodynamic characteristics of the nacelle that was tested, increasing the fan pressure ratio to approximately 1.35 would have resulted in a maximum propulsive efficiency of 67 percent.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-3064 , E-7824
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  • 58
    Publication Date: 2019-06-27
    Description: A research program was conducted to define the level of the current technology base for reaction control system rocket engines suitable for space shuttle applications. The project consisted of engine analyses, design, fabrication, and tests. The specific objectives are: (1) extrapolating current engine design experience to design of an RCS engine with required safety, reliability, performance, and operational capability, (2) demonstration of multiple reuse capability, and (3) identification of current design and technology deficiencies and critical areas for future effort.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134336 , REPT-8701-910041
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  • 59
    Publication Date: 2019-06-27
    Description: A technique and apparatus for optically detonating insensitive high explosives, is disclosed. An explosive device is formed by containing high explosive material in a house having a transparent window. A thin metallic film is provided on the interior surface of the window and maintained in contact with the high explosive. A laser pulse provided by a Q-switched laser is focussed on the window to vaporize the metallic film and thereby create a shock wave which detonates the high explosive. Explosive devices may be concurrently or sequentially detonated by employing a fiber optic bundle to transmit the laser pulse to each of the several individual explosive devices.
    Keywords: PROPULSION SYSTEMS
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  • 60
    Publication Date: 2019-06-27
    Description: A single-stage axial-flow turbine was investigated to determine the effect of increased stator throat area on the performance level of a turbojet engine turbine. The stator blades were bent to increase the throat area in order to move the compressor operating point in the engine farther away from surge. Results are compared with those obtained with the as-cast stator setting.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-7626 , E-7776
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  • 61
    facet.materialart.
    Unknown
    In:  CASI
    Publication Date: 2019-06-27
    Description: The JT8D turbofan engine, widely used in short and medium range transport aircraft, contributes substantially to airport community noise. The jet noise is predominant in the JT8D engine and may be reduced in a modified engine, without loss of thrust, by increasing the airflow to reduce jet velocity. A configuration study evaluated the effects of fan airflow, fan pressure ratio, and bypass ratio on noise, thrust, and fuel comsumption. The cycle selected for the modified engine was based upon an increased diameter, single-stage fan and two additional core engine compressor stages, which replace the existing two-stage fan. Modifications were also made to the low pressure turbine to provide the increased torque required by the larger diameter fan. The resultant JT8D-100 engine models have the following characteristics at take-off thrust, compared to the current JT8D engine: Airflow and bypass ratio are increased, and fan pressure ratio and engine speed are reduced. The resultant engine is also longer, larger in diameter, and heavier than the JT8D base model, but these latter changes are compensated by the increased thrust and decreased fuel comsumption of the modified engine, thus providing the capability for maintaining the performance of the current JT8D-powered aircraft.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134654 , PWA-4790
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  • 62
    Publication Date: 2019-06-27
    Description: At a typical STOL aircraft takeoff and landing velocity, wind tunnel aerodynamic and acoustic measurements demonstrated that an inlet lip-area contraction ratio of 1.35 was superior to a ratio of 1.26 at high incidence angles. A 17 percent reduction in net thrust and an increase of 9 decibels in sound pressure level at the blade passing frequency resulted from inlet flow separation at an incidence angle of 50 deg with the 1.26-contraction-ratio inlet. Reverse-thrust forces obtained with blade rotation through the feathered angle were 1.8 times larger than with blade rotation through the flat angle. Reverse-thrust force was reduced from 30 to 50 percent and sound pressure level increased from 3 to 7 decibels at the blade passing frequency between the wind-tunnel-off condition and a typical STOL aircraft landing velocity.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-3062 , E-7844
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  • 63
    Publication Date: 2019-06-27
    Description: The acoustic investigation and evaluation of the C propulsion turbofan engine are discussed. The engine was built as a part of the Quiet Engine Program. The objectives of the program are as follows: (1) to determine the noise levels produced turbofan bypass engines, (2) to demonstrate the technology and innovations which will reduce the production and radiation of noise in turbofan engines, and (3) to acquire experimental acoustic and aerodynamic data for high bypass turbofan engines to provide a better understanding of noise production mechanisms. The goals of the program called for a turbofan engine 15 to 20 PNdB quieter than currently available engines in the same thrust class.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-121176 , R73AEG364
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  • 64
    Publication Date: 2019-06-27
    Description: Experimental results of a series of engine tests designed to obtain the stage dynamics of an eight-stage axial-flow compressor over the frequency range of 0.5 to 200 hertz are presented. The total pressure at the compressor face was varied by means of a secondary air jet system installed in the engine inlet and positioned to oppose the primary airflow. Total-pressure probes located at each compressor stage were used to obtain the frequency response of each compressor-stage total pressure to the average compressor-inlet total pressure. The engine operating conditions were chosen to illustrate the effects of changing the rotor speed, changing the exhaust nozzle area, and isolating the compressor discharge pressure perturbations from the fuel control and hence, the fuel flow.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-3012 , E-7709
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  • 65
    Publication Date: 2019-06-27
    Description: The design and evaluation are described of a digital turbojet engine control which is capable of sensing catastrophic failures in either the engine rotor speed or the compressor discharge static-pressure signal and is capable of switching control modes to maintain near normal operation. The control program was developed for and tested on a turbojet engine located in a sea-level test stand. The control program is also capable of acquiring all the data that are necessary for the fail-operational control to function.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-3043 , E-7761
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  • 66
    Publication Date: 2019-06-27
    Description: A 24.70-cm(9.72-in.) tip diameter single-stage, axial-flow turbine was designed for a turbojet engine. Cold-air tests were made over a range of speeds from 30 to 105 percent of design equivalent speed and over a range of total-pressure ratios from 1.30 to 2.66. Results are presented in terms of specific work, torque, mass flow, and efficiency.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-7625 , E-7754
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  • 67
    Publication Date: 2019-06-27
    Description: A series of investigations was conducted episodically from the 1950's to the early 1970's to investigate the feasibility and potential problem areas in the use of hydrogen fuel for gas turbine engines. A brief summary and bibliography are presented of the research that has been conducted by NASA, its predecessor NACA, and by industry under U. S. Air Force sponsorship. Although development efforts would be required to provide hydrogen fueled gas turbine engines for aircraft, past research has shown that hydrogen fueled engines are feasible, and except for flight weight liquid hydrogen pumps, there are no problem areas relating to engines requiring significant research.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-71561 , E-7990
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  • 68
    Publication Date: 2019-06-27
    Description: Two-F-111A airplanes were subjected to conditions that caused engine compressor surges and accompanying duct hammershock pressure transients. Flight speed ranged from Mach 0.71 to Mach 2.23, and altitude varied from approximately 3200 meters to 14,500 meters. A wide range of compressor pressure ratios was covered. Stabilized free-stream, engine, and duct conditions were established before each compressor surge. Dynamic pressure instrumentation at the compressor face and in the duct recorded the pressure transients associated with the surges. Hammershock pressures were analyzed with respect to the stabilized conditions preceding the compressor surges. The hammershock transients caused large pressure rises at the compressor face and in the duct. Hammershock pressure ratios at the compressor face were not affected by free-stream Mach number or altitude but were functions of engine variables, such as compressor pressure ratio. The maximum hammershock pressure ratio of approximately 1.83 occurred at a compressor pressure ratio of approximately 21.7.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-7696 , H-804
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  • 69
    facet.materialart.
    Unknown
    In:  CASI
    Publication Date: 2019-06-27
    Description: A study was conducted to determine the promising propulsion systems for advanced supersonic transport application, and to identify the critical propulsion technology requirements. It is shown that noise constraints have a major effect on the selection of the various engine types and cycle parameters. Several promising advanced propulsion systems were identified which show the potential of achieving lower levels of sideline jet noise than the first generation supersonic transport systems. The non-afterburning turbojet engine, utilizing a very high level of jet suppression, shows the potential to achieve FAR 36 noise level. The duct-heating turbofan with a low level of jet suppression is the most attractive engine for noise levels from FAR 36 to FAR 36 minus 5 EPNdb, and some series/parallel variable cycle engines show the potential of achieving noise levels down to FAR 36 minus 10 EPNdb with moderate additional penalty. The study also shows that an advanced supersonic commercial transport would benefit appreciably from advanced propulsion technology. The critical propulsion technology needed for a viable supersonic propulsion system, and the required specific propulsion technology programs are outlined.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134633 , PWA-TM-4871
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  • 70
    Publication Date: 2019-06-27
    Description: The overall and blade-element performances and the aerodynamic design parameters of a transonic rotor with an aspect ratio of 6.5, designed to investigate the effects of aspect ratio on range and performance are presented. The rotor was designed for a total-pressure ratio of 1.53, an efficiency of 0.898, and a weight flow of 67.76 pounds per second.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-7662 , E-7641
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  • 71
    Publication Date: 2019-06-27
    Description: Test hardware, facilities, and procedures are described along with results of electrically heated tube and channel tests conducted to determine adverse operating condition limits for convectively cooled chambers typical of Space Shuttle Orbit Manuevering Engine designs. Hot-start tests were conducted with corrosion resistant steel and nickel tubes with both monomethylhydrazine and 50-50 coolants. Helium ingestion, in both bubble and froth form, was studied in tubular test sections. Helium bubble ingestion and burn-out limits in rectangular channels were also investigated.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134282 , ASR-74-131
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  • 72
    Publication Date: 2019-06-27
    Description: The overall and blade-element performance of two configurations of a moderately high aspect ratio transonic compressor rotor are presented. The subject rotor has conventional blade dampers. The performance is compared with a rotor utilizing dual wire friction dampers. At design speed the subject achieved a pressure ratio of 1.52 and efficiency of 0.89 at a near design weight flow of 72.1 pounds per second. The rotor with wire dampers gave consistently higher pressure ratios at each speed, but efficiencies for the two rotors were about the same. Stall margin for the subject rotor was 20.4 percent, but for the wire damped rotor only 4.0 percent.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-3041 , E-7068
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  • 73
    Publication Date: 2019-06-27
    Description: 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.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134286 , R-9270
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  • 74
    Publication Date: 2019-06-27
    Description: Two aspects of the dispersion of pollutants from aircraft are reviewed. The first is the dispersal of aircraft exhaust emissions in the vicinity of airports; the second is the dispersal of exhaust trails in the upper atmosphere. Techniques available for modeling this dispersal and how they might be applied to the airport problem are discussed. Field studies of airport pollution are then reviewed to assess current pollutant levels around airports and the aircraft's contribution to those levels. The possibility of contrail formation from jet emissions at high altitude is then considered and the effect of uncertainties in the trial mixing processes evaluated.
    Keywords: PROPULSION SYSTEMS
    Type: AGARD The Fluid Dynamics Aspects of Air Pollution Related to Aircraft Operations; p 5-16
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  • 75
    Publication Date: 2019-06-27
    Description: Arcs representative of those occurring between the grids of a mercury ion thruster were simulated. Parameters affecting an arc and the resulting damage were studied. The parameters investigated were arc energy, arc duration, and grid geometry. Arc attenuation techniques were also investigated. Potentially serious damage occurred at all energy levels representative of actual thruster operating conditions. Of the grids tested, the lowest open-area configuration sustained the least damage for given conditions. At a fixed energy level a long duration discharge caused greater damage than a short discharge. Attenuation of arc current using various impedances proved to be effective in reducing arc damage. Faults were also deliberately caused using chips of sputtered materials formed during the operation of an actual thruster. These faults were cleared with no serious grid damage resulting using the principles and methods developed in this study.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-3044 , E-7767
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  • 76
    Publication Date: 2019-06-27
    Description: Tests were conducted on a high-tip-speed low-loading transonic fan stage to determine the performance and inlet flow distortion tolerance of the design. Test data were recorded for overall and blade element performance with both uniform and distorted inlet flows. A tabular summary of the data and a representative selection of the computer data reduction sheets are presented.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-121262 , AIRESEARCH-73-9487
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  • 77
    Publication Date: 2019-06-27
    Description: Tests were conducted on a high-tip-speed, low-loading transonic fan stage to determine the performance and inlet flow distortion tolerance of the design. The fan was designed for high efficiency at a moderate pressure ratio by designing the hub section to operate at minimum loss when the tip operates with an oblique shock. The design objective was an efficiency of 86 percent at a pressure ratio of 1.5, a specific flow (flow per unit annulus area) of 42 lb/sec-sq. ft (205.1 kgm/sec-m sq), and a tip speed of 1600 ft/sec (488.6 m/sec). During testing, a peak efficiency of 84 percent was achieved at design speed and design specific flow. At the design speed and pressure ratio, the flow was 4 percent greater than design, efficiency was 81 percent, and a stall margin of 24 percent was obtained. The stall line was improved with hub radial distortion but was reduced when the stage was tested with tip radial and circumferential flow distortions. Blade-to-blade values of static pressures were measured over the rotor blade tips.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-121263 , AIRESEARCH-73-9488
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  • 78
    Publication Date: 2019-06-27
    Description: A low-power MPD thruster with downstream cathode was tested for endurance with a series of hollow cathode designs. Failure modes and failure mechanisms were identified. A new hollow cathode (with rod inserts) has emerged which shows promise for long life. The downstream positioning of the cathode was also changed from an on-axis location to an off-axis location. Data are presented for a 1332-hour life test of this new hollow cathode located at the new off-axis location. Xenon propellant was used.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-7616 , E-7647
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  • 79
    Publication Date: 2019-06-27
    Description: An analytical procedure for computing the performance of supersonic ejector nozzles is presented. This procedure includes real sonic line effects and an interaction analysis for the mixing process between the two streams. The procedure is programmed in FORTRAN 4 and has operated successfully on IBM 7094, IBM 360, CDC 6600, and Univac 1108.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-7602 , E-7476
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  • 80
    Publication Date: 2019-06-27
    Description: The results of a program of experimental and analytical research in casing treatments over axial compressor rotor blade tips are presented. Circumferential groove, axial-skewed slot, and blade angle slot treatments were tested. These yielded, for reduction in stalling flow and loss in peak efficiency, 5.8% and 0 points, 15.3% and 2.0 points, and 15.0% and 1.2 points, respectively. These values are consistent with other experience. The favorable stalling flow situations correlated well with observations of higher-then-normal surface pressures on the rotor blade pressure surfaces in the tip region, and with increased maximum diffusions on the suction surfaces. Annular wall pressure gradients, especially in the 50-75% chord region, are also increased and blade surface pressure loadings are shifted toward the trailing edge for treated configurations. Rotor blade wakes may be somewhat thinner in the presence of good treatments, particularly under operating conditions close to the baseline stall.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134552 , R73AEG326
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  • 81
    Publication Date: 2019-06-27
    Description: Methods of handling one-third-octave band noise data originating from the outdoor full-scale fan noise facility and the engine acoustic facility at the Lewis Research Center are presented. Procedures for standardizing, retrieving, extrapolating, and reporting these data are explained. Computer programs are given which are used to accomplish these and other noise data analysis tasks. This information is useful as background for interpretation of data from these facilities appearing in NASA reports and can aid data exchange by promoting standardization.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-3013 , E-7609
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  • 82
    Publication Date: 2019-06-27
    Description: A ten-year, life-cycle cost comparison is made of the following engines for small helicopter use: (1) simple turboshaft; (2) regenerative turboshaft; (3) compression-ignition reciprocator; (4) spark-ignited rotary; and (5) spark-ignited reciprocator. Based on a simplified analysis and somewhat approximate data, the simple turboshaft engine apparently has the lowest costs for mission times up to just under 2 hours. At 2 hours and above, the regenerative turboshaft appears promising. The reciprocating and rotary engines are less attractive, requiring from 10 percent to 80 percent more aircraft to have the same total payload capability as a given number of turbine powered craft. A nomogram was developed for estimating total costs of engines not covered in this study.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-71517 , E-7908
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  • 83
    Publication Date: 2019-06-27
    Description: Propulsion system tests were conducted on a full scale Centaur vehicle to investigate system capability of the proposed D-lT configuration for a three-burn mission. This particular mission profile requires that the engines be capable of restarting and firing for a final maneuver after a 5-1/2-hour coast to synchronous orbit. The thermal conditioning requirements of the engine and propellant feed system components for engine start under these conditions were investigated. Performance data were also obtained on the D-lT type computer controlled propellant tank pressurization system. The test results demonstrated that the RL-10 engines on the Centaur vehicle could be started and run reliably after being thermally conditioned to predicted engine start conditions for a one, two and three burn mission. Investigation of the thermal margins also indicated that engine starts could be accomplished at the maximum predicted component temperature conditions with prestart durations less than planned for flight.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-71511 , E-7889
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  • 84
    Publication Date: 2019-06-27
    Description: Experimental results of a three-stage highly loaded fan drive turbine test program are presented. A plain blade turbine, a tandem blade turbine, and a tangentially leaned stator turbine were designed for the same velocity diagram and flowpath. Seven combinations of bladerows were tested to evaluate stage performances and effects of the tandem blading and leaned stator. The plain blade turbine design point total-to-total efficiency was 0.886. The turbine with the stage three leaned stator had the same efficiency with an improved exit swirl profile and increased hub reaction. Two-stage group tests showed that the two-stage turbine with tandem stage two stator had an efficiency of 0.880 compared to 0.868 for the plain blade two-stage turbine.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-2362 , R73AEG154
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  • 85
    Publication Date: 2019-06-27
    Description: Similarity parameters from the equations of motion of particles immersed in a gas flow are derived. These parameters relate the particles which follow a certain trajectory in an equivalent cold gas turbine to particles that will follow the same trajectory in a real hot gas turbine. Numerical solutions of the trajectories that particles follow in the vortex and rotor regions of a radial inflow turbine are used to verify the range of Reynolds numbers in which the derived similarity parameters are applicable. In addition, an example is presented of typical particle sizes that can be observed in high speed photographic data collection and at the same time simulate the trajectories of particles in a real hot gas turbine.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134589
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  • 86
    Publication Date: 2019-07-20
    Description: A comprehensive summary of the results of a cold-flow and hot-fire experimental study of the mixing and atomization characteristics of injector elements incorporating noncircular orifices is presented. 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.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134315 , R-9271
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  • 87
    Publication Date: 2019-07-13
    Description: Optimal solid-rocket thrust profiles for the parallel-burn, solid-rocket-assisted space shuttle are investigated. Solid-rocket thrust profiles are simulated by using third-degree spline functions, with the values of the thrust ordinates defined as parameters. The profiles are optimized parametrically, using the Davidon-Fletcher-Powell penalty function method, by minimizing propellant weight subject to state and control inequality constraints and to terminal boundary conditions. This study shows that optimizing a control variable parametrically by using third-degree spline function interpolation allows the control to be shaped so that inequality constraints are strictly adhered to and all corners are eliminated. The absence of corners, which is realistic in nature, makes this method attractive from the viewpoint of solid rocket grain design.
    Keywords: PROPULSION SYSTEMS
    Type: AIAA PAPER 74-823 , Mechanics and Control of Flight Conference; Aug 05, 1974 - Aug 09, 1974; Anaheim, CA
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  • 88
    Publication Date: 2019-07-13
    Description: Description of a simplified one-dimensional treatment of fuel injection for supersonic combustor performance analysis. Representative mixing efficiency variations for both parallel and cross-stream injection are obtained, and approximate means for estimating the length required for complete mixing are demonstrated. Comparisons of calculated and measured data show good agreement.
    Keywords: PROPULSION SYSTEMS
    Type: International Symposium on Air Breathing Engines; Mar 24, 1974 - Mar 29, 1974; Sheffield
    Format: text
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  • 89
    Publication Date: 2019-07-13
    Description: In order to effectively use a compressor face total-pressure distortion index as a measure of inlet-engine compatibility, a correlation of distortion amplitude with stall margin must be developed with minimal scatter. A recent analysis of data recorded in extensive distortion screen tests with the J85-GE-13 turbojet engine has resulted in a correlation based on compressor discharge pressure ratioed to the minimum pressure at the compressor face. Simply by determining compressor stall lines with a single hub radial distortion pattern, a single tip radial pattern, and with undistorted inflow, the overall compressor pressure ratio at stall for even the most complex distortion pattern was found to be predictable.
    Keywords: PROPULSION SYSTEMS
    Type: National Conference on Environmental Effects on Aircraft and Propulsion Systems; May 21, 1974 - May 23, 1974; Trenton, NJ
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  • 90
    facet.materialart.
    Unknown
    In:  Other Sources
    Publication Date: 2019-07-13
    Description: Hydrogen is injected from a downstream facing step in a wall into a high-temperature stream. Temperature and hydroxyl radical concentration are measured downstream of the injection plane by means of ultraviolet absorption spectroscopy. The experimental results are compared with theory which is based on a finite-difference solution of boundary-layer equations. Finite-rate kinetics equations are included in the analysis. The analytic predictions are also compared with previously obtained experimental results which are based on probe measurements. Comparison is made between calculated and observed ignition distances.
    Keywords: PROPULSION SYSTEMS
    Type: Joint Fluids Engineering Conference; May 13, 1974 - May 15, 1974; Montreal; Canada
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  • 91
    Publication Date: 2019-07-13
    Description: The program discussed is aimed at improving performance within an extended operating range, reducing weight, increasing service life, achieving greater cost effectiveness, reducing noise and exhaust pollution, and improving means of energy conservation. The program places emphasis on basic research in numerous related technical disciplines, system analysis, component technology, full scale propulsion system studies, and technology demonstrations. Much attention in the discussion is given to noise and pollution minimization research.
    Keywords: PROPULSION SYSTEMS
    Type: ICAS PAPER 74-51 , International Council of the Aeronautical Sciences Congress; Aug 25, 1974 - Aug 30, 1974; Haifa; Israel
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  • 92
    Publication Date: 2019-07-13
    Description: Flyover and static noise data from several engines are presented that show inlet fan noise measured in flight can be lower than that projected from static tests for some engines. The differences between flight and static measurements appear greatest when the fan fundamental tone due to rotor-stator interaction or to the rotor-alone field is below cutoff. Data from engine and fan tests involving inlet treatment on the walls only are presented that show the attenuation from this treatment is substantially larger than expected from previous theories or flow duct experience. Data showing noise shielding effects due to the location of the engine on the airplane are also presented. These observations suggest that multiringed inlets may not be necessary to achieve the desired noise reduction in many applications.
    Keywords: PROPULSION SYSTEMS
    Type: AIAA PAPER 74-946 , Aircraft Design, Flight Test and Operations Meeting; Aug 12, 1974 - Aug 14, 1974; Los Angeles, CA
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  • 93
    Publication Date: 2019-07-13
    Description: Aspects of the development and characteristics of thermal shock resistant hafnia ceramic material for use in solid propellant rocket nozzles are presented. The investigation of thermal shock resistance factors for hafnia based composites, and the preparation and analysis of a model of elastic materials containing more than one crack are reported.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-138804 , QPR-12
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  • 94
    facet.materialart.
    Unknown
    In:  CASI
    Publication Date: 2019-07-13
    Description: The results of impact tests on large, fiber composite fan blades for aircraft turbofan engine applications are discussed. Solid composite blades of two different sizes and designs were tested. Both graphite/epoxy and boron/epoxy were evaluated. In addition, a spar-shell blade design was tested that had a boron/epoxy shell bonded to a titanium spar. All blades were tested one at a time in a rotating arm rig to simulate engine operating conditions. Impacting media included small gravel, two inch diameter ice balls, gelatin and RTV foam-simulated birds, as well as starlings and pigeons. The results showed little difference in performance between the graphite and boron/epoxy blades. The results also indicate that composite blades may be able to tolerate ice ball and small bird impacts but need improvement to tolerate birds in the small duck and larger category.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-71544 , Natl. SAMPE Symp. and Exhibition; Apr 23, 1974 - Apr 25, 1974; Buena Park, CA; United States
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  • 95
    Publication Date: 2019-07-13
    Description: The objectives, program plan, schedule, pollution and performance goals, program approaches to pollution reduction, and status of the project are presented. Advanced supersonic transport and combustion noise programs are also discussed. The phases of the program consist of the following: (1) screening of low pollutant combustors, (2) refinement of the best combustors, and (3) engine demonstration of the best combustors. It is stated that all combustor designs investigated appear capable of achieving the idle pollutant goals with additional development effort.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-71547 , E-7956 , Air Transport Meeting; Apr 30, 1974 - May 02, 1974; Dallas
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  • 96
    Publication Date: 2019-07-13
    Description: A summary is presented of an acoustic test program for investigating engine noise suppression and jet/flap interaction noise associated with an EBF STOL powered lift system. A highly suppressed TF-34 engine and EBF wing were used in the investigation. The engine was suppressed 21 PndB to a level of 94 PndB. An UTW powered lift system was tested with conventional, mixer, and decayer-type nozzles. The configuration with velocity decayer nozzle and acoustically treated shroud had the lowest noise (98 PndB). An OTW configuration with non-decayer nozzle was about 10 db quieter than the corresponding UTW system. UTW and OTW noise data are compared with scale model correlations.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-71539 , E-7946 , Air Transport Meeting; Apr 30, 1974 - May 02, 1974; Dallas
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  • 97
    Publication Date: 2019-07-13
    Description: Projected minimum levels of engine exhaust emissions that may be practicably achievable for future commercial aircraft operating at high altitude cruise conditions are presented. The forecasts are based on: (1) current knowledge of emission characteristics of combustors and augmentors; (2) the current status of combustion research in emission reduction technology; (3) predictable trends in combustion systems and operating conditions as required for projected engine designs that are candidates for advanced subsonic or supersonic commercial aircraft. Results are presented for cruise conditions in terms of an emission index, g pollutant/kg fuel. Two sets of engine exhaust emission predictions are presented: the first, based on an independent NASA study and the second, based on the consensus of an ad hoc committee composed of industry, university, and government representatives. The consensus forecasts are in general agreement with the NASA forecasts.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-71509 , E-7822 , Conf. on Climatic Impact Assessment Program; Feb 26, 1974 - Mar 01, 1974; Cambridge, MA; United States
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  • 98
    Publication Date: 2019-07-13
    Description: Results from experimental work on model scale cascade reversers with cold airflow are presented. Sound power level directivity and spectral characteristics for cascade reversers are reported. Effect of cascade exit area ratio, vane profile shape, and emission arc are discussed. Model equivalent diameters varied from 3 to 5 inches, pressure ratios range from 1.15 to 3.0. Depending on the reverser type, acoustic power was proportional to the 4 1/2 to 6th power of ideal jet velocity. Reverser noise peaked at higher frequency and was more omnidirectional than nozzle-alone jet noise. Appreciable reduction in sideline noise was obtained from plane shields. Airfoil-vaned cascades were the most aerodynamically efficient and least noisy reversers. Scaling of cascade reverser data to example aircraft engines showed all cascades above the 95 PNdB sideline goal from STOL aircraft. However, the airfoil-vaned reverser has a good potential for meeting this goal for high-bypass (low pressure ratio) exhausts.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-71500 , E-7864 , Aerospace Sci. Meeting; Jan 30, 1974 - Feb 01, 1974; Washington, DC; United States
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  • 99
    Publication Date: 2019-07-13
    Description: A review is presented of recent experimental results, analytical procedures and test techniques employed to evaluate the effects of inlet flow distortion on the stability characteristics of representative afterburning turbofan and turbojet compression systems. Circumferential distortions of pressure and temperature, separately and in combination are considered. Resulting engine sensitivity measurements are compared with predictions based on simplified parallel compressor models and with several distortion descriptor parameters.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-71505 , E-7878 , Aerospace Sci. Meeting; Jan 30, 1974 - Feb 01, 1974; Washington, DC; United States
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  • 100
    facet.materialart.
    Unknown
    In:  CASI
    Publication Date: 2019-07-13
    Description: The aerodynamic drag and fuel distribution patterns of injectors designed for a supersonic combustion ramjet were measured at Mach numbers of 2, 2.5, and 3. The most significant parameter effecting the drag was found to be the injector thickness ratio. A two-fold reduction in the thickness ratio caused a 65 percent decrease in drag. Changing the injector sweep angle a factor of 2 resulted in only a small change in drag. A reversal of injector sweep, from sweepback to sweepforward, did not change the measured drag. Helium gas was injected through the struts to simulate the penetration and spreading patterns of hydrogen. Sampling measurements were made at approximately 2 duct heights downstream of the combustor. The spacing required between fuel injectors was found to be about 10 jet diameters. The effect of gas injection on the measured drag was found to be minor.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-71504 , E-7849 , Symp. on the Fluid Mech. of Combustion; May 13, 1974 - May 16, 1974; Montreal
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