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At least 631 records · Page 35

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NASA NTRS1973

Pressure signatures for a .00053 scale model of the Saturn 5-Apollo launch vehicle with simulated exhaust plumes

Wind tunnel pressure signatures are presented for 10 Mach numbers over a range from 3.01 to 7.29 for a .00053-scale model of the Saturn 5 Apollo launch vehicle complete with escape tower and solid body simulated exhaust plumes for each Mach number. The effect of simulated plume length on the wind tunnel pressure signature was investigated at Mach 4.01. An analysis of the error incurred by extrapolating pressure signatures having strong shock waves by weak shock procedures is presented.

Hicks, R. M.↗

NASA NTRS1973

Dispersion and dilution of jet aircraft exhaust at high-altitude flight conditions

A method is presented for estimating the dispersion and dilution of jet aircraft exhaust from aircraft passage through times on the order of weeks thereafter. In the near wake of the aircraft, the solution is that for round turbulent jets in a parallel flow. More rapid dispersion due to atmospheric effects begins when the scale-dependent eddy viscosity becomes larger than the turbulent jet eddy viscosity. In the far wake region, the solution approaches that for scale-dependent dispersion from a point source moving with the aircraft. Calculations are presented for supersonic aircraft at high altitude flight conditions.

Holdeman, J. D.↗

NASA NTRS1973

Effects of prevaporized fuel on exhaust emissions of an experimental gas turbine combustor.

Effect of fuel vaporization on the exhaust emission levels of oxides of nitrogen (NOx), carbon monoxide, total hydrocarbons and smoke number were obtained in an experimental turbojet combustor segment. Two fuel injector types were used in which liquid ASTM A-1 jet fuel and vapor propane fuel were independently controlled to simulate varying degrees of vaporization. Tests were conducted over a range of inlet-air temperatures from 478 to 700 K, pressures from 4 to 20 atmospheres, and combustor reference velocities from 15.3 to 27.4 m/sec. Converting from liquid to complete vapor fuel resulted in NOx reductions as much as 22% and smoke number reductions up to 51%.

Norgren, C. T.↗

NASA NTRS1972

Analysis and design of three dimensional supersonic nozzles. Volume 1: Nozzle-exhaust flow field analysis by a reference plane characteristics technique

A second order numerical method employing reference plane characteristics has been developed for the calculation of geometrically complex three dimensional nozzle-exhaust flow fields, heretofore uncalculable by existing methods. The nozzles may have irregular cross sections with swept throats and may be stacked in modules using the vehicle undersurface for additional expansion. The nozzles may have highly nonuniform entrance conditions, the medium considered being an equilibrium hydrogen-air mixture. The program calculates and carries along the underexpansion shock and contact as discrete discontinuity surfaces, for a nonuniform vehicle external flow.

Dash, S.↗

NASA NTRS1972

Analysis and design of three dimensional supersonic nozzles. Volume 2: Numerical program for analysis of nozzle-exhaust flow fields

The FORTRAN IV Program developed to analyze the flow field associated with scramjet exhaust systems is presented. The instructions for preparing input and interpreting output are described. The program analyzes steady three dimensional supersonic flow by the reference plane characteristic technique. The governing equations and numerical techniques employed are presented in Volume 1 of this report.

Kalben, P.↗

NASA NTRS1973

Analysis of stratified and closely spaced jets exhausting into a crossflow

Procedures have been developed for determining the flow field about jets with velocity stratification exhausting into a crossflow. Jets with three different types of exit velocity stratification have been considered: (1) jets with a relatively high velocity core; (2) jets with a relatively low velocity core; and (3) jets originating from a vaned nozzle. The procedure developed for a jet originating from a high velocity core nozzle is to construct an equivalent nozzle having the same mass flow and thrust but having a uniform exit velocity profile. Calculations of the jet centerline and induced surface static pressures have been shown to be in good agreement with test data for a high velocity core nozzle. The equivalent ideal nozzle has also been shown to be a good representation for jets with a relatively low velocity core and for jets originating from a vaned nozzle in evaluating jet-induced flow fields. For the singular case of a low velocity core nozzle, namely a nozzle with a dead air core, and for the vaned nozzle, an alternative procedure has been developed. The internal mixing which takes place in the jet core has been properly accounted for in the equations of motion governing the jet development. Calculations of jet centerlines and induced surface static pressures show good agreement with test data these nozzles.

Ziegler, H.↗

NASA NTRS1974

Jet engine exhaust emissions of high altitude commercial aircraft projected to 1990

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 status of combustion research in emission reduction technology; and (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 fueled by either JP fuel, liquefied natural gas, or hydrogen. Results are presented for cruise conditions in terms of both an emission index (g constituent/kg fuel) and an emission rate (g constituent/hr).

Grobman, J.↗

NASA NTRS1974

Acoustic loads on an externally blown flap system due to impingement of a TF-34 jet engine exhaust

One of the powered lift systems currently being considered for use on STOL vehicles obtains additional lift by impinging the flow of a jet exhaust on the flaps and turning it downward. This direct impingement subjects large areas of the flap to the fluctuating pressure loads of the turbulent jet flow; these loads have the potential for causing high vibration levels and sonic fatigue failures. A test was conducted on a full-scale mockup of an externally blown flap system, which consisted of a TF-34 engine and a section of a wing having a triple-slotted flap. Data were obtained and evaluated on the fluctuating pressures on the flaps and it is the purpose of this paper to present some of the results describing the loads.

Schoenster, J. A.↗

NASA NTRS1974

Interference between exhaust system and afterbody of twin-engine fuselage configurations

Some of the mutual aircraft afterbody and engine nozzle interferences that can exist on aircraft complex aft-ends and on simplified twin-jet afterbodies are reviewed. Emphasis is placed on studies of twin-engine fuselage configurations with nozzles installed near the terminus of the afterbody where the interactions of the nozzle exhausts and external stream produce a complex flow-field environment. The magnitude of the aft-end problem for real aircraft configurations is illustrated, and some theoretical and empirical considerations of afterbody drag on simple bodies are treated. Many of the factors regarding airframe-installation effects on nozzle performance are discussed, as well as nozzle-installation effects on aircraft performance.

Runckel, J. F.↗

NASA NTRS1974

Gaseous exhaust emissions from a J-58 engine at simulated supersonic flight conditions

Emissions of total oxides of nitrogen, unburned hydrocarbons, carbon monoxide, and carbon dioxide from a J-58 engine at simulated flight conditions of Mach 2.0, 2.4, and 2.8 at 19.8 km altitude are reported. For each flight condition, measurements were made for four engine power levels from maximum power without afterburning through maximum afterburning. These measurements were made 7 cm downstream of the engine primary nozzle using a single point traversing gas sample probe. Results show that emissions vary with flight speed, engine power level, and with radial position across the exhaust.

Holdeman, J. D.↗

NASA NTRS1974

Effects of equivalence ratio and dwell time on exhaust emissions from an experimental premixing prevaporizing burner

A flame-tube study was performed to determine the effects of equivalence ratio and residence time on exhaust emissions with premixed, prevaporized propane fuel. Nitrogen oxides emissions as low as .3 g NO2/kg fuel were measured with greater than 99% combustion efficiency at 800 K inlet temperature and an equivalence ratio of .4. For a constant combustion efficiency, lower nitrogen oxides emissions were obtained by burning very lean with relatively long residence times than by using somewhat higher equivalence ratios with shorter times.

Anderson, D.↗

NASA NTRS1973

Exhaust pressure and density of various pulsed MPD-Arc thruster systems

Exhaust flow in a new 155-cm-i.d. vacuum facility is compared with earlier measurements in a small (15.2-cm-i.d.) duct. Reductions in post-transient impact pressure are about 5:1 in the larger facility. Corresponding reduced electron number densities (about 2 x 10 to the 13th power per cu cm) are noted. A new 125-microsec pulse-forming network power source produced no major differences in impact pressure compared to the crowbarred condenser bank used earlier. Comparing a puff gas feed of the arc chamber with a new 10-msec steady gas feed also shows no major difference in impact pressure for 125-microsec powering.

Michels, C. J.↗

NASA NTRS1974

Dispersion and dilution of jet aircraft exhaust at high-altitude flight conditions

Holdeman, J. D.↗

NASA NTRS1974

Ingestion and dispersion of engine exhaust products by trailing vortices for supersonic flight in the stratosphere

Detailed analysis has been made of the ingestion and dispersion of engine exhaust products into the trailing-vortex system of a supersonic aircraft flying in the stratosphere. The rate of mixing between the supersonic jet and the coflowing supersonic stream was found to be an order of magnitude less than would be expected on the basis of subsonic eddy-viscosity results. Ingestion started at the end of the potential core, and all hot gas from the engine was ingested into the trailing vortex within two core lengths. The temperature rise in the wake of the supersonic aircraft was found to be much greater than that for a subsonic transport, since temperature rise varies directly as speed squared and inversely as aspect ratio. As a result, the final buoyancy phase started only ten seconds after leaving the trailing edge. Comparison between the buoyancy calculations for the supersonic case with nondimensionalized subsonic aircraft contrail data on wake spreading showed good agreement.

Nielsen, J. N.↗

NASA NTRS1952

Analysis of a Pneumatic Probe for Measuring Exhaust-Gas Temperatures with Some Preliminary Experimental Results

A pneumatic probe based on continuity of mass flow through two restrictions separated by a cooling chamber was constructed to measure gas temperature at and beyond the limit of thermocouples. This probe consisted of a subsonic flat-plate orifice for the first restriction and a sonic-flow converging-diverging nozzle for the second restriction. The effect of variation in gas constants on the calibration is examined for common engine-exhaust gases. A high-temperature wind tunnel that allowed calibration of the probe at temperatures up to 2000 deg R and. Mach numbers up to 0.8 is described. Agreement to better than 30 deg R between pneumatic probe indication and the indication of a rake of radiation shielded thermocouples indicates that extrapolation of the calibration to higher temperatures is possible with fair accuracy

Scadron, Marvin D.↗

NASA NTRS1974

Optical properties of mercury ion thruster exhausts and implications for science instruments

Emission from the exhaust plume of a 30 cm mercury ion thruster was measured from 160 to 600 nm as a function of axial and radial distance from the thruster discharge chamber. The spectrally dispersed absolute intensities were used to construct an empirical volume rate function. The function was integrated along a typical instrument field of view, and the resulting apparent brightness was compared with instrument sensitivities to evaluate the extent of optical interference. Most of the emitted radiation came from UV lines of excited mercury atoms and ions, with no observable continuum emission. The intensity levels degraded rapidly with distance from the thruster so that optical interference was negligible for fields of view not intercepting the beam axis. The operation of only one instrument, a zodiacal photopolarimeter, was considered incompatible with simultaneous thruster operation.

Monahan, K. M.↗

NASA NTRS1974

Exhaust emissions of a double annular combustor: Parametric study

A full scale double-annular ram-induction combustor designed for Mach 3.0 cruise operation was tested. Emissions of oxides of nitrogen, carbon monoxide, unburned hydrocarbons, and smoke were measured over a range of combustor operating variables including reference velocity, inlet air temperature and pressure, and exit average temperature. ASTM Jet-A fuel was used for these tests. An equation is provided relating oxides of nitrogen emissions as a function of the combustor, operating variables. A small effect of radial fuel staging on reducing exhaust emissions (which were originally quite low) is demonstrated.

Schultz, D. F.↗

NASA NTRS1975

External fins and ejector action for reducing the infrared emission of engine exhaust ducting

An analytical investigation was conducted to determine the feasibility of using external fins and ejector action on the exhaust ducting of a helicopter to reduce the infrared emission of the aircraft. Temperatures were calculated for both circular disk fins and pin fins. Results show that combining ejector action with fins can lower the metal temperature to acceptable levels at least for high flight speeds.

Vanfossen, G. J., Jr.↗