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At least 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS2004

Exhaust Nozzle for a Multitube Detonative Combustion Engine

An improved type of exhaust nozzle has been invented to help optimize the performances of multitube detonative combustion engines. The invention is applicable to both air-breathing and rocket engines used to propel some aircraft and spacecraft, respectively. In a detonative combustion engine, thrust is generated through the expulsion of combustion products from a detonation process in which combustion takes place in a reaction zone coupled to a shock wave. The combustion releases energy to sustain the shock wave, while the shock wave enhances the combustion in the reaction zone. The coupled shockwave/reaction zone, commonly referred to as a detonation, propagates through the reactants at very high speed . typically of the order of several thousands of feet per second (of the order of 1 km/s). The very high speed of the detonation forces combustion to occur very rapidly, thereby contributing to high thermodynamic efficiency. A detonative combustion engine of the type to which the present invention applies includes multiple parallel cylindrical combustion tubes, each closed at the front end and open at the rear end. Each tube is filled with a fuel/oxidizer mixture, and then a detonation wave is initiated at the closed end. The wave propagates rapidly through the fuel/oxidizer mixture, producing very high pressure due to the rapid combustion. The high pressure acting on the closed end of the tube contributes to forward thrust. When the detonation wave reaches the open end of the tube, it produces a blast wave, behind which the high-pressure combustion products are expelled from the tube. The process of filling each combustion tube with a detonable fuel/oxidizer mixture and then producing a detonation repeated rapidly to obtain repeated pulses of thrust. Moreover, the multiple combustion tubes are filled and fired in a repeating sequence. Hence, the pressure at the outlet of each combustion tube varies cyclically. A nozzle of the present invention channels the expansion of the pulsed combustion gases from the multiple combustion tubes into a common exhaust stream, in such a manner as to enhance performance in two ways: (1) It reduces the cyclic variations of pressure at the outlets of the combustion tubes so as to keep the pressure approximately constant near the optimum level needed for filling the tubes, regardless of atmospheric pressure at the altitude of operation; and (2) It maximizes the transfer of momentum from the exhaust gas to the engine, thereby maximizing thrust. The figure depicts a typical engine equipped with a nozzle according to the invention. The nozzle includes an interface section comprising multiple intake ports that couple the outlets of the combustion tubes to a common plenum. Proceeding from its upstream to its downstream end, the interface section tapers to a larger cross-sectional area for flow. This taper fosters expansion of the exhaust gases flowing from the outlets of the combustion tubes and contributes to the desired equalization of exhaust combustion pressure. The cross-sectional area for flow in the common plenum is greater than, or at least equal to, the combined cross-sectional flow areas of the combustor tubes. In the common plenum, the exhaust streams from the individual combustion tubes mix to form a single compound subsonic exhaust stream. Downstream of the common plenum is the throat that tapers to a smaller flow cross section. In this throat, the exhaust gases become compressed to form a compound sonic gas stream. Downstream of the throat is an expansion section, which typically has a bell or a conical shape. (The expansion section can be truncated or even eliminated in the case of an air-breathing engine.) After entering the expansion section, the exhaust gases expand rapidly from compound sonic to compound supersonic speeds and are then vented to the environment. The basic invention admits of numerous variations. For example, the combustion tubes can be arranged around the central axin a symmetrical or asymmetrical pattern other than the one shown in the figure. For another example, the flow cross-sectional area(s) of one or more of the intake ports in the interface section, of the common plenum, the throat, and/or the expansion section can be varied, either symmetrically or asymmetrically, to adjust dynamics of the exhaust stream or to direct the thrust vector away from the central axis.

Bratkovich, Thomas E.↗
NASA NTRS2024

Multiple Subscale Magnetic Reconnection Embedded inside a Heliospheric Current Sheet Reconnection Exhaust: Evidence for Flux Rope Merging

We report observations of multiple subscale reconnecting current sheets embedded inside a large-scale heliospheric current sheet (HCS) reconnection exhaust. The discovery was made possible by the unusual skimming trajectory of Parker Solar Probe through a sunward-directed HCS exhaust, sampling structures convecting with the exhaust outflows for more than 3 hr during Encounter 14, at a radial distance of ∼17 solar radii. A large number of subscale current sheets (SCSs) were detected inside the HCS exhaust. Remarkably, five SCSs showed direct evidence for reconnection, displaying near-Alfvénic outflow jets and bifurcated current sheets. The reconnecting SCSs all had small magnetic shears (27°–81°), i.e., strong guide fields. The thickness of the subscale reconnecting current sheets ranged from ∼60 km to ∼5000 km (∼20–2000 ion inertial lengths). The SCS exhausts were directed predominantly in the normal or out-of-plane direction of the HCS, i.e., nearly orthogonal to the HCS exhaust direction. The presence of multiple low-magnetic-shear reconnecting current sheets inside a large-scale exhaust could be associated with coalescence of multiple large flux ropes inside the HCS exhaust. The orientation of some SCS exhausts was partly in the ecliptic plane of the HCS, which may indicate that the coalescence process is highly three-dimensional. Since the coalescence process is likely short-lived, the detection of five such events inside a single HCS crossing could imply the common occurrence of flux rope coalescence in large-scale HCS reconnection exhausts.

T D Phan↗
NASA NTRS1947

Effect of Exhaust Pressure on the Cooling Characteristics of a Liquid-Cooled Engine

Data for a liquid-cooled engine with a displacement volume of 1710 cubic inches were analyzed to determine the effect of exhaust pressure on the engine cooling characteristics. The data covered a range of exhaust pressures from 7 to 62 inches of mercury absolute, inlet-manifold pressures from 30 to 50 inches of mercury absolute, engine speeds from 1600 to 3000 rpm, and fuel-air ratios from 0.063 to 0.100. The effect of exhaust pressure on engine cooling was satisfactorily incorporated in the NACA cooling-correlation method as a variation in effective gas temperature with exhaust pressure. Large variations of cylinder-head temperature with exhaust pressure were obtained for operation at constant charge flow. At a constant charge flow of 2 pounds per second (approximately 1000 bhp) and a fuel-air ratio of 0.085, an increase in exhaust pressure from 10 to 60 inches of mercury absolute resulted in an increase of 40 F in average cylinder-head temperature. For operation at constant engine speed and inlet-manifold pressure and variable exhaust pressure (variable charge flow), however, the effect of exhaust pressure on cylinder-head temperature is small. For example, at an inlet-manifold pressure of 40 inches of mercury absolute, an engine speed of 2400 rpm.- and a fuel-air ratio of 0.085, the average cylinder-head temperature was about the same at exhaust pressures of 10 and 60 inches of,mercury absolute; a rise and a subsequent decrease of about 70 occurred between these extremes.

Doyle, Ronald B.↗
NASA NTRS2021

On the Presence and Thermalization of Cold Ions in the Exhaust of Antiparallel Symmetric Reconnection

Using fully kinetic 2.5 dimensional particle-in-cell simulations of anti-parallel symmetric magnetic reconnection, we investigate how initially cold ions are captured by the reconnection process, and how they evolve and behave in the exhaust. We find that initially cold ions can remain cold deep inside the exhaust. Cold ions that enter the exhaust downstream of active separatrices, closer to the dipolarization front, appear as cold counter-streaming beams behind the front. In the off-equatorial region, these cold ions generate ion-acoustic waves that aid in the thermalization both of the incoming and outgoing populations. Closest to the front, due to the stronger magnetization, the ions can remain relatively cold during the neutral plane crossing. In the intermediate exhaust, the weaker magnetization leads to enhanced pitch angle scattering and reflection. Cold ions that enter the exhaust closer to the X line, at active separatrices, evolve into a thermalized exhaust. Here, the cold populations are heated through a combination of thermalization at the separatrices and pitch angle scattering in the curved magnetic field around the neutral plane. Depending on where the ions enter the exhaust, and how long time they have spent there, they are accelerated to different energies. The superposition of separately thermalized ion populations that have been accelerated to different energies form the hot exhaust population.

Cecilia Norgren↗
NASA NTRS1974

Effluent sampling of Scout D and Delta launch vehicle exhausts

Characterization of engine-exhaust effluents (hydrogen chloride, aluminum oxide, carbon dioxide, and carbon monoxide) has been attempted by conducting field experiments monitoring the exhaust cloud from a Scout-Algol III vehicle launch and a Delta-Thor vehicle launch. The exhaust cloud particulate size number distribution (total number of particles as a function of particle diameter), mass loading, morphology, and elemental composition have been determined within limitations. The gaseous species in the exhaust cloud have been identified. In addition to the ground-based measurements, instrumented aircraft flights through the low-altitude, stabilized-exhaust cloud provided measurements which identified CO and HCI gases and Al2O3 particles. Measurements of the initial exhaust cloud during formation and downwind at several distances have established sampling techniques which will be used for experimental verification of model predictions of effluent dispersion and fallout from exhaust clouds.

Hulten, W. C.↗
NASA NTRS1975

An experimental study of jet exhaust simulation

Afterbody drag predictions for jet aircraft are usually made experimentally with the jet exhaust flow simulated. The physical gas properties of the fluid used for the model jet exhaust can affect the accuracy of simulation of the airplane's jet exhaust plume. The effect of the accuracy of this simulation on afterbody drag was investigated by wind-tunnel tests with single engine model. In addition to unheated air as the exhaust gas, the decomposition products of three different concentrations of hydrogen peroxide were utilized. The air jet simulation consistently resulted in higher boattail drag than hydrogen peroxide simulation. The differences in drag for the various exhaust gases are attributed to different plume shapes and entrainment properties of the gases. The largest differences in drag due to exhaust gas properties were obtained for the combination of high transonic Mach numbers and high boattail angles. For these conditions, the current data indicate that the use of air to simulate a nonafterburning turbojet exhaust can result in an increase in afterbody amounting to 20 percent of the nonafterburning turbojet value.

Compton, W. B., III↗
NASA NTRS1979

Concepts for reducing exhaust emissions and fuel consumption of the aircraft piston engine

A study was made to reduce exhaust emissions and fuel consumption of a general aviation aircraft piston engine by applying known technology. Fourteen promising concepts such as stratified charge combustion chambers, cooling cylinder head improvements, and ignition system changes were evaluated for emission reduction and cost effectiveness. A combination of three concepts, improved fuel injection system, improved cylinder head with exhaust port liners and exhaust air injection was projected as the most cost effective and safe means of meeting the EPA standards for CO, HC and NO. The fuel economy improvement of 4.6% over a typical single engine aircraft flight profile does not though justify the added cost of the three concepts, and significant reductions in fuel consumption must be applied to the cruise mode where most of the fuel is used. The use of exhaust air injection in combination with exhaust port liners reduces exhaust valve stem temperatures which can result in longer valve guide life. The use of exhaust port liners alone can reduce engine cooling air requirements by 11% which is the equivalent of a 1.5% increase in propulsive power. The EPA standards for CO, HC and NO can be met in the IO-520 engine using air injection alone or the Simmonds improved fuel injection system.

Rezy, B. J.↗
NASA NTRS1979

Argon/UF6 plasma exhaust gas reconstitution experiments using preheated fluorine and on-line diagnostics

The feasibility of employing a flowing, high-temperature, pure fluorine/UF6 regeneration system to efficiently convert a large fraction of the effluent plasma exhaust back to pure UF6 was demonstrated. The custom built T.O.F. mass spectrometer sampling system permitted on-line measurements of the UF6 concentration at different locations in the exhaust system. Negligible amounts ( 100 ppm) of UF6 were detected in the axial bypass exhaust duct and the exhaust ducts downstream of the cryogenic trap system used to collect the UF6, thus verifying the overall system efficiency over a range of operating conditions. Use of a porous Monel duct as part of the exhaust duct system, including provision for injection of pure fluorine, provided a viable technique to eliminate uranium compound residue on the inside surface of the exhaust ducts. Typical uranium compound mass deposition per unit area of duct was 2 micron g/sq cm. This porous duct technique is directly applicable to future uranium compound transfer exhaust systems. Throughout these experiments, additional basic data on the corrosion aspects of hot, pressurized UF6/fluorine were also accumulated.

Roman, W. C.↗
NASA NTRS1993

Solid propellant exhausted aluminum oxide and hydrogen chloride - Environmental considerations

Measurements of gaseous hydrogen chloride (HCl) and particulate aluminum oxide (Al2O3) were made during penetrations of five Space Shuttle exhaust clouds and one static ground test firing of a shuttle booster. Instrumented aircraft were used to penetrate exhaust clouds and to measure and/or collect samples of exhaust for subsequent analyses. The focus was on the primary solid rocket motor exhaust products, HCl and Al2O3, from the Space Shuttle's solid boosters. Time-dependent behavior of HCl was determined for the exhaust clouds. Composition, morphology, surface chemistry, and particle size distributions were determined for the exhausted Al2O3. Results determined for the exhaust cloud from the static test firing were complicated by having large amounts of entrained alkaline ground debris (soil) in the lofted cloud. The entrained debris may have contributed to neutralization of in-cloud HCl.

Cofer, W. R., III↗
NASA NTRS1931

A Method for Reducing the Temperature of Exhaust Manifolds

This report describes tests conducted at the Langley Memorial Aeronautical Laboratory on an "air-inducting" exhaust manifold for aircraft engines. The exhaust gases from each cylinder port are discharged into the throat of an exhaust pipe which has a frontal bellmouth. Cooling air is drawn into the pipe, where it surrounds and mixes with the exhaust gases. Temperatures of the manifold shell and of the exhaust gases were obtained in flight for both a conventional manifold and the air-inducting manifold. The air-inducting manifold was installed on an engine which was placed on a test stand. Different fuels were sprayed on and into the manifold to determine whether the use of this manifold reduced the fire hazard. The flight tests showed reductions in manifold temperatures of several hundred degrees, to values below the ignition point of aviation gasoline. On the test stand when the engine was run at idling speeds fuels sprayed into the manifold ignited. It is believed that at low engine speeds the fuel remained in the manifold long enough to become thoroughly heated, and was then ignited by the exhaust gas which had not mixed with cooling air. The use of the air-inducting exhaust manifold must reduce the fire hazard by virtue of its lower operating temperature, but it is not a completely satisfactory solution of the problem.

Schey, Oscar W↗
NASA NTRS1974

Wind tunnel investigation of simulated helicopter engine exhaust interacting with windstream

A wind tunnel investigation of the windstream-engine exhaust flow interaction on a light observation helicopter model has been conducted in the Langley V/STOL tunnel. The investigation utilized flow visualization techniques to determine the cause to determine the cause of exhaust shield overheating during cruise and to find a means of eliminating the problem. Exhaust flow attachment to the exhaust shield during cruise was found to cause the overheating. Several flow-altering devices were evaluated to find a suitable way to correct the problem. A flow deflector located on the model cowling upstream of the exhaust in addition to aerodynamic shield fairings provided the best solution. Also evaluated was heat transfer concept employing pin fins to cool future exhaust hardware. The primary flow visualization technique used in the investigation was a newly developed system employing neutrally buoyant helium-filled bubbles. The resultant flow patterns were recorded on motion picture film and on television magnetic tape.

Shaw, C. S.↗
NASA NTRS1981

Prediction of jet exhaust noise on airframe surfaces during flight

The behavior of pressure fluctuations measured on the airframe of a prototype high lift jet transport (YC-14) is presented. The data are characterized in terms of a particular jet exhaust flow field idealization, jet mixing noise, and exhaust shock noise. Generalized spectrum shapes and scaling relations for peak level and frequency of peak level are developed, and the frequency is found to depend on jet exhaust velocity and aircraft velocity. Comparisons are made with near-field engine exhaust noise of a conventional jet, and results suggest that the same two exhaust noises are important for both aircraft types. Surface fluctuating pressure data are assessed, and results suggest that the jet mixing and exhaust shock noise source characterizations for the YC-14 have useful applicability to conventionally configured jets.

Butzel, L. M.↗
NASA NTRS1988

The effect of exhaust plume/afterbody interaction on installed Scramjet performance

Newly emerging aerospace technology points to the feasibility of sustained hypersonic flight. Designing a propulsion system capable of generating the necessary thrust is now the major obstacle. First-generation vehicles will be driven by air-breathing scramjet (supersonic combustion ramjet) engines. Because of engine size limitations, the exhaust gas leaving the nozzle will be highly underexpanded. Consequently, a significant amount of thrust and lift can be extracted by allowing the exhaust gases to expand along the underbody of the vehicle. Predicting how these forces influence overall vehicle thrust, lift, and moment is essential to a successful design. This work represents an important first step toward that objective. The UWIN code, an upwind, implicit Navier-Stokes computer program, has been applied to hypersonic exhaust plume/afterbody flow fields. The capability to solve entire vehicle geometries at hypersonic speeds, including an interacting exhaust plume, has been demonstrated for the first time. Comparison of the numerical results with available experimental data shows good agreement in all cases investigated. For moderately underexpanded jets, afterbody forces were found to vary linearly with the nozzle exit pressure, and increasing the exit pressure produced additional nose-down pitching moment. Coupling a species continuity equation to the UWIN code enabled calculations indicating that exhaust gases with low isentropic exponents (gamma) contribute larger afterbody forces than high-gamma exhaust gases. Moderately underexpanded jets, which remain attached to unswept afterbodies, underwent streamwise separation on upswept afterbodies. Highly underexpanded jets produced altogether different flow patterns, however. The highly underexpanded jet creates a strong plume shock, and the interaction of this shock with the afterbody was found to produce complicated patterns of crossflow separation. Finally, the effect of thrust vectoring on vehicle balance has been shown to alter dramatically the vehicle pitching moment.

Edwards, Thomas Alan↗
NASA NTRS1991

A parametric experimental investigation of a scramjet nozzle at Mach 6 with Freon and argon or air used for exhaust simulation

A parametric experimental investigation of a scramjet nozzle was conducted with a gas mixture used to simulate the scramjet engine exhaust flow at a free-stream Reynolds number of approximately 6.5 x 10(exp 6) per foot. External nozzle surface angles of 16, 20, and 24 deg were tested with a fixed-length ramp and for cowl internal surface angles of 6 and 12 deg. Pressure data on the external nozzle surface were obtained for mixtures of Freon and argon gases with a ratio of specific heats of about 1.23, which matches that of a scramjet exhaust. Forces and moments were determined by integration of the pressure data. Two nozzle configurations were also tested with air used to simulate the exhaust flow. On the external nozzle surface, lift and thrust forces for air exhaust simulation were approximately half of those for Freon-argon exhaust simulation and the pitching moment was approximately a third. These differences were primarily due to the difference in the ratios of specific heats between the two exhaust simulation gases. A 20 deg external surface angle produced the greatest thrust for a 6 deg cowl internal surface angle. A flow fence significantly increased lift and thrust forces over those for the nozzle without a flow fence.

Cubbage, James M.↗
NASA NTRS1990

Analysis of exhaustive limited service for token ring networks

Token ring operation is well-understood in the cases of exhaustive, gated, gated limited, and ordinary cyclic service. There is no current data, however, on queueing models for the exhaustive limited service type. This service type differs from the others in that there is a preset maximum (omega) on the number of packets which may be transmitted per token reception, and packets which arrive after token reception may still be transmitted if the preset packet limit has not been reached. Exhaustive limited service is important since it closely approximates a timed token service discipline (the approximation becomes exact if packet lengths are constant). A method for deriving the z-transforms of the distributions of the number of packets present at both token departure and token arrival for a system using exhaustive limited service is presented. This allows for the derivation of a formula for mean queueing delay and queue lengths. The method is theoretically applicable to any omega. Fortunately, as the value of omega becomes large (typically values on the order of omega = 8 are considered large), the exhaustive limited service discipline closely approximates an exhaustive service discipline.

Peden, Jeffery H.↗
NASA NTRS2004

Exhaust Simulation Testing of a Hypersonic Airbreathing Model at Transonic Speeds

An experimental study was performed to examine jet-effects for an airframe-integrated, scramjet-rocket combined-cycle vehicle configuration at transonic test conditions. This investigation was performed by testing an existing exhaust simulation wind tunnel model, known as Model 5B, in the NASA Langley 16-Ft. Transonic Tunnel. Tests were conducted at freestream Mach numbers from 0.7 to 1.2, at angles of attack from 2 to +14 degrees, and at up to seven nozzle static pressure ratio values for a set of horizontal-tail and body-flap deflections. The model aftbody, horizontal tails, and body flaps were extensively pressure instrumented to provide an understanding of jet-effects and control-surface/plume interactions, as well as for the development of analytical methodologies and calibration of computational fluid dynamic codes to predict this type of flow phenomenon. At all transonic test conditions examined, the exhaust flow at the exit of the internal nozzle was over-expanded, generating an exhaust plume that turned toward the aftbody. Pressure contour plots for the aftbody of Model 5B are presented for freestream transonic Mach numbers of 0.70, 0.95, and 1.20. These pressure data, along with shadowgraph images, indicated the impingement of an internal plume shock and at least one reflected shock onto the aftbody for all transonic conditions tested. These results also provided evidence of the highly three-dimensional nature of the aftbody exhaust flowfield. Parametric testing showed that angle-of-attack, static nozzle pressure ratio, and freestream Mach number all affected the exhaust-plume size, exhaust-flowfield shock structure, and the aftbody-pressure distribution, with Mach number having the largest effect. Integration of the aftbody pressure data showed large variations in the pitching moment throughout the transonic regime.

Huebner, Lawrence D.↗
NASA NTRS2002

Real-time exhaust gas modular flowmeter and emissions reporting system for mobile apparatus

A real-time emissions reporting system includes an instrument module adapted to be detachably connected to the exhaust pipe of a combustion engine to provide for flow of exhaust gas therethrough. The instrument module includes a differential pressure probe which allows for determination of flow rate of the exhaust gas and a gas sampling tube for continuously feeding a sample of the exhaust gas to a gas analyzer or a mounting location for a non-sampling gas analyzer. In addition to the module, the emissions reporting system also includes an elastomeric boot for detachably connecting the module to the exhaust pipe of the combustion engine, a gas analyzer for receiving and analyzing gases sampled within the module and a computer for calculating pollutant mass flow rates based on concentrations detected by the gas analyzer and the detected flowrate of the exhaust gas. The system may also include a particulate matter detector with a second gas sampling tube feeding same mounted within the instrument module.

Breton, Leo Alphonse Gerard↗
NASA NTRS2009

Shape memory alloy actuated adaptive exhaust nozzle for jet engine

The proposed adaptive exhaust nozzle features an innovative use of the shape memory alloy (SMA) actuators for actively control of the opening area of the exhaust nozzle for jet engines. The SMA actuators remotely control the opening area of the exhaust nozzle through a set of mechanism. An important advantage of using SMA actuators is the reduction of weight of the actuator system for variable area exhaust nozzle. Another advantage is that the SMA actuator can be activated using the heat from the exhaust and eliminate the need of other energy source. A prototype has been designed and fabricated. The functionality of the proposed SMA actuated adaptive exhaust nozzle is verified in the open-loop tests.

Song, Gangbing↗