Search NASA⌕ Search

SEARCH · Search NASA

Results for “AIR INLET”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 289 records · Page 16
All sourcesNASA NTRSDOE OSTI
DOE OSTI2021

Prediction of Detonation-Induced Disturbances Propagating Upstream into Inlets of Rotating Detonation Combustors

Disturbances caused by the detonation wave in a rotating detonation combustor (RDC) propagate upstream through the inlet, and can potentially affect and couple to upstream components, such as turbomachinery or isolators. These disturbances can potentially also affect the operation of the RDC itself. By drawing from the analogy of a detonation wave bounded by an inert gas, the pressure disturbances observed upstream of the inlet are explained as the consequence of the passage of an upstream propagating oblique shock. In this study, the pressure rise in the plenum from the oblique shock is measured in an axial air inlet RDC. The speed of the upstream propagating wave is estimated to be moderately above the acoustic speed of the oxidizer in the plenum. The wave propagates into the plenum despite local regions of choking in the inlet. It is estimated that the time it takes a fluid particle to transit from the plenum to the detonation channel through the inlet is much larger than the rotational time of the detonation wave. This implies that a fluid particle experiences multiple shocks prior to entering the detonation channel. The oblique shock propagating upstream through the inlet area change is modeled by leveraging an analogy with a quasi-1D shock wave moving in a variable area duct with mean (incoming) flow. Due to flow expansion along the area change, fluid particles are found to experience stronger shocks in the inlet than in the plenum, thereby creating different thermodynamic states within the fill region as the oxidizer emerges from the inlet.

Feleo, Alexander↗
NASA NTRS1947

Performance of the 19XB 10-Stage Axial-Flow Compressor

The 19xB compressor, which replaces the 19B coaapreseor and has the same length and diameter 88 the 19B compressor, was designed with 10 stages to deliver 30 pounds of air per second for a pressure ratio of 4.17 at an equivalent speed of 17,000 rpm; the 19B was designed with six stages for a pressure ratio of 2.7 at the same weight flow and speed as the 19XB compressor. The performance characteristics of the new compressor were determined at the NACA Cleveland laboratory at the request of the Bureau of Aeronautics, Navy Department. Results are presented of the investigation made to evaluate the over-all performance of the compressor, the effects of possible leakage past the rotor rear air seal, the effects of inserting instruments in each row of stator blades and in the first row of outlet guide vanes, and the effects of changing the temperature and the pressure of the inlet air. The results of the interstage surveys are also presented.

Downing, Richard M.↗
NASA NTRS1976

Cold-air annular-cascade investigation of aerodynamic performance of core-engine-cooled turbine vanes. 2: Pressure surface trailing edge ejection and split trailing edge ejection

The aerodynamic performance of two trailing edge ejection cooling configurations of a core-engine stator vane were experimentally determined in an ambient inlet-air full-annular cascade where three-dimensional effects could be obtained. The tests were conducted at the design mean-radius ideal aftermixed critical velocity ratio of 0.778. Overall vane aftermixed thermodynamic and primary efficiencies were obtained over a range of coolant flows to about 10 percent of the primary flow at a primary to coolant total temperature ratio of 1.0. The radial variation in efficiency and the circumferential and radial variations in vane-exit total pressure were determined. Comparisons are made with the solid (uncooled) vane.

Mclallin, K. L.↗
DOE OSTI2023

Technoeconomic assessment of distillation heat transfer intensification using aeroelastically fluttering reeds

Dry cooling, where forced air is the heat transfer medium, is a preferable cooling method in arid locations lacking readily available process water. However, such locations often experience high ambient temperatures that limit the effectiveness of air cooling. The objective of this study is to quantify the economic and energetic benefits of heat transfer intensification via the implementation of aeroelastically fluttering reeds to the air-cooled condenser of a methanol distillation column. Condenser size and performance, regarding recovered methanol and required fan power, is evaluated across condenser operating temperatures (T cond ) from 60 to 62°C and heat transfer coefficients (U) U base –2U base for a range of inlet air temperatures based on ambient temperature data from Yuma, Arizona. Under typical design sizing, condenser capital cost was reduced by 6%–35% (1.3U base –2U base ) and nominal methanol recovery was increased from 0.26% to 0.38% (T cond = 62–60°C). At optimized condenser size, all enhanced U and T cond pairs increase methanol recovery and reduce fan power costs compared to the optimal Ubase reference. Overall, using enhanced heat transfer to maintain condenser temperature under a wider range of inlet conditions, rather than to reduce operation temperature, produces more favorable performance. Methanol price is not a determining factor in which pairs are profitable. Analysis was repeated for a global warming scenario, revealing more valuable improvements under elevated temperatures. Energy savings from condenser improvement to a methanol production system are not significant with respect to an optimized conventional system. Furthermore, unit economic and energetic incentives suggest implementation of fluttering reeds may be justified in other applications.

42 ENGINEERING↗
DOE OSTI2023

Experimental Study of the Invariance of Pressure Gain with Respect to the Dynamics of Multiple Competing Waves in a Rotating Detonation Combustor

Changes in the overall performance of a rotating detonation combustor with respect to changes in operation mode and wave dynamics arising by operation with fixed inlet/exit geometry but at different combustor, lengths are investigated experimentally. The air inlet, fuel injection, and exit constriction geometry are held constant while only the length of the detonation channel is varied from 71 to 137 mm (which corresponds to about 10 to 20 channel widths). Operation of H2/air over a range of air mass flow rates and equivalence ratios are considered for every chamber length. The number and speed of (primary) detonation and secondary waves are characterized through high-speed pressure measurements in the detonation channel and aft chemiluminescence videos. The number of waves is found to increase with length while detonation wave speed decreases significantly. Particular emphasis is given to characterize a phenomenon that is observed at operation with longer combustor lengths and higher mass flow rates. The phenomenon manifests as a super-cycle behavior with a period equal to many detonation wave rotational periods and is characterized by a periodic and structured ascending/descending sequence of the number, speed, and direction of both (primary) detonation and secondary waves. This phenomenon is likely a manifestation of the system failing in achieving operation with a higher number of detonation waves as length and/or mass flow rate are increased. The performance of the device is quantified in terms of measured thrust and pressure gain (through the use of the equivalent available pressure). Both metrics are essentially found to be invariant with respect to combustor length and most importantly, mode of operation. Surprisingly, even operation with complex wave dynamics arising from transitions between multiple competing wave systems does not appear to alter the overall global performance of the device but rather, it remains defined by the total (capture) air mass flow rate, equivalence ratio, and inlet/outlet areas.

33 ADVANCED PROPULSION SYSTEMS↗
NASA NTRS1984

An interactive computer code for calculation of gas-phase chemical equilibrium (EQLBRM)

A user friendly, menu driven, interactive computer program known as EQLBRM which calculates the adiabatic equilibrium temperature and product composition resulting from the combustion of hydrocarbon fuels with air, at specified constant pressure and enthalpy is discussed. The program is developed primarily as an instructional tool to be run on small computers to allow the user to economically and efficiency explore the effects of varying fuel type, air/fuel ratio, inlet air and/or fuel temperature, and operating pressure on the performance of continuous combustion devices such as gas turbine combustors, Stirling engine burners, and power generation furnaces.

Pratt, B. S.↗
NASA NTRS1958

A Preliminary Investigation of the Performance of a Short-length Turbojet Combustor Using Vaporized Hydrocarbon Fuels

Two short turbojet combustors designed for use with vaporized hydrocarbon fuels were tested in a one-quarter annular duct. The experimental combustors consisted of many small "swirl-can" combustor elements manifolded together. This design approach allowed the secondary mixing zone to be considerably reduced over that of conventional combustors. The over-all combustion lengths, for the two configurations were 13.5 and 11.0 inches, approximately one-half the length of the shortest conventional combustors. These short combustors did not provide combustion efficiencies as high as those for conventional combustors at low pressures. However, over the range of combustor-inlet total-pressures expected in aircraft capable of flight at Mach numbers of 2.5 and above, these short combustors gave very high efficiencies. A combustion efficiency of 97 percent was obtained at a combustor-inlet total-pressure of 25.0 inches of mercury absolute, reference velocity of 120 feet per second, and inlet-air total temperature of 1160 deg R. By proportioning the fuel flow between the manifold rows of can combustor elements, control of the combustor-outlet radial total-temperature profile was demonstrated. Combustor totalpressure loss varied from 0.75 percent of the inlet total pressure at isothermal conditions and a reference velocity of 75 feet per second to 5.5 percent at a total-temperature ratio of 1.8 and a reference velocity of 180 feet per second.

Jones, R. E.↗
NASA NTRS1979

Atomization of water jets and sheets in axial and swirling airflows

Axial and swirling airflows were used to break up water jets and sheets into sprays of droplets to determine the overall effects of orifice diameter, weight flow of air, and the use of an air swirler on fineness of atomization as characterized by mean drop size. A scanning radiometer was used to determine the mean drop diameter of each spray. Swirling airflows were produced with an axial combustor, 70 deg blake angle, air swirling. Water jets were injected axially upstream, axially downstream and cross stream into the airflow. In addition, pressure atomizing fuel nozzles which produced a sheet and ligament type of breakup were investigated. Increasing the weight flow rate of air or the use of an air swirling markedly reduced the spray mean drop size. Test conditions included a water flow rate of 68.0 liter per hour and airflow rates (per unit area) of 3.7 to 25.7 g per square cm per sec, at 293 K and inlet-air static pressures of 1.01 x 10 to the 5th to 1.98 x 10 to the 5th N/sq m.

Ingebo, R. D.↗
NASA NTRS1984

Heat pipes to reduce engine exhaust emissions

A fuel combustor is presented that consists of an elongated casing with an air inlet conduit portion at one end, and having an opposite exit end. An elongated heat pipe is mounted longitudinally in the casing and is offset from and extends alongside the combustion space. The heat pipe is in heat transmitting relationship with the air intake conduit for heating incoming air. A guide conduit structure is provided for conveying the heated air from the intake conduit into the combustion space. A fuel discharge nozzle is provided to inject fuel into the combustion space. A fuel conduit from a fuel supply source has a portion engaged in heat transfer relationship of the heat pipe for preheating the fuel. The downstream end of the heat pipe is in heat transfer relationship with the casing and is located adjacent to the downstream end of the combustion space. The offset position of the heat pipe relative to the combustion space minimizes the quenching effect of the heat pipe on the gaseous products of combustion, as well as reducing coking of the fuel on the heat pipe, thereby improving the efficiency of the combustor.

Schultz, D. F.↗
NASA NTRS1979

Airborne atmospheric sampling system

The atmospheric sampling system developed for use on board commercial airliners as part of the Global Atmospheric Sampling Program (GASP) is described. The automated air-constituent measuring system is installed in a Boeing 747 airliner below the passenger cabin floor near the nose wheel well. It consists of an air sample flow system, composed of air inlet and pressurization systems, computerized data acquisition and system control units which direct system operation in 15 modes, and commercial instruments significantly modified to measure low levels of atmospheric constituents (ozone, water vapor, nitrogen oxides, carbon monoxide, chlorofluoromethanes, particulates, condensation nuclei, sulfates and nitrates). Flight and meteorological data, including air temperature and altitude, are also recorded. The system is designed for servicing at 14-day intervals, and to require a minimum of aircrew involvement.

Gustafsson, U. R. C.↗
NASA NTRS1984

Flame radiation

Total radiation and heat flux data was obtained on a combustor liner by advanced instrumentation. If the results obtained by the special instrumentation are considered to be representative of the total radiation and heat flux, then the effect of variation of engine operating parameters and of fuel type can be more easily obtained. The special instrumentation used for these investigations consisted of five total radiometers and two total heat flux gages. The radiometers were arranged axially and circumferentially through sliding air seals in the outer liner. The two heat flux gages were welded in the outer liner between two circumferential radiometers. Static pressures were obtained on both the cold and the hot side of the outer liner in the area of the heat flux gages. Liner metal temperatures were also obtained. The combustor inlet pressure was varied over a nominal range of 0.5 to 2.07 MPa, inlet air temperature from 550 to 670 K, and fuel-air ratio from about 0.015 to 0.040. The two fuels tested were ASTM Jet A and a fuel designated as ERBS V. Results of the tests are discussed.

Wear, J. D.↗
NASA NTRS1976

Reduction of gaseous pollutant emissions from gas turbine combustors using hydrogen-enriched jet fuel

Recent progress in an evaluation of the applicability of the hydrogen enrichment concept to achieve ultralow gaseous pollutant emission from gas turbine combustion systems is described. The target emission indexes for the program are 1.0 for oxides of nitrogen and carbon monoxide, and 0.5 for unburned hydrocarbons. The basic concept utilizes premixed molecular hydrogen, conventional jet fuel, and air to depress the lean flammability limit of the mixed fuel. This is shown to permit very lean combustion with its low NOx production while simulataneously providing an increased flame stability margin with which to maintain low CO and HC emission. Experimental emission characteristics and selected analytical results are presented for a cylindrical research combustor designed for operation with inlet-air state conditions typical for a 30:1 compression ratio, high bypass ratio, turbofan commercial engine.

Clayton, R. M.↗
NASA NTRS1975

Combustion intensity and distribution relation to noise generation

Experiments with several different flame holder geometries were conducted to investigate the degree to which combustion roughness can be altered by altering the flame intensity and flame distribution in a ducted combustion system. The effect of admitting primary air through a plane-slotted or a slotted-swirl vane flame holder was compared and the combustion roughness and noise was contrasted with that obtained with a closed front-end perforated can. The slotted front-end burners produced much smoother burning and less noise than the closed front-end can. No advantage was apparent with swirl vs nonswirl when approximately the same inlet flow distribution was maintained. Preheated inlet air provided somewhat smoother combustion as compared with ambient temperature air. The combustion roughness with methyl alcohol was briefly compared with that of isooctane; indications are that it burns more smoothly, but more detailed studies are needed to substantiate these indications.

Plett, E. G.↗
NASA NTRS1947

Flight Investigation of the Effects of Ice on an I-16 Jet-Propulsion Engine

A flight investigation of an I-16 jet propulsion engine installed in the waist compartment of a B-24M airplane was made to determine the effect of induction-system icing on the performance of the engine. Flights were made at inlet-air temperatures of 15 deg, 20 deg., and 25 F, an indicated airspeed of 180 miles per hour, jet-engine speeds of 13,000 and 15,000 rpm, liquid-water contents of approximately 0.3 to 0.5 gram per cubic meter, and an average water droplet size of approximately 50 microns. Under the most severe icing conditions obtained, ice formed on the screen over the front inlet to the compressor and obstructed about 70 percent of the front-inlet area. The thrust was thereby reduced 13.5 percent, the specific fuel consumption increased 17 percent, and the tail-pipe temperature increased 82 F. No icing of the rear compressor-inlet screen was encountered.

Pragliola, Philip C.↗
NASA NTRSDate not supplied

Design Concepts for an Optimized, Modular, and Scalable Liquid Amine Test Stand

The Air Revitalization Labs at NASA Ames Research Center (ARC) is investigating liquid amines to for their performance removing spacecraft carbon dioxide (CO2) for crewed space missions. Development of the current one-crew scale amine test stand stems from several modifications to a test stand initially built at NASA Johnson Space Center (JSC) This sub-scale system has been modified to incorporate a flow spreader for the inlet air stream, an updated contactor design for CO2 capture, a degasser to regenerate the liquid sorbent and recover the captured CO2, and a condenser to reduce water loss. Results from the existing integrated liquid amine test stand have provided valuable insight, but continued performance optimization requires design and build of a new integrated test system. New design concepts will institute a modular configuration to allow for interfacing with novel contactor and degasser architecture, parallel or series flow pathways for the air stream and liquid sorbent, and scaling up to a 4-crew CO2 removal system.

Tiago Faria Costa↗
NASA NTRS1981

Effects of fuel-injector design on ultra-lean combustion performance

Emissions data were obtained for six fuel injector configurations tested with ultra lean combustion. Fuel injectors included three multiple source designs and three configurations using a single air assist injector. Only the multiple source fuel injectors provided acceptable emissions. Values of 16g CO/kg fuel, 1.9g HC/kg fuel, and 19.g NO2/kg fuel were obtained for the combustion temperature range of 1450 to 1700 K for both a high blockage 19 source injector and a low blockage 41 source injector. It was shown that high fuel injector pressure drop may not be required to achieve low emissions performance at high inlet air temperature when the fuel is well dispersed in the airstream.

Anderson, D. N.↗
NASA NTRS1956

Low-speed cascade investigation of loaded leading-edge compressor blades

Six percent thick NACA 63-series compressor-blade sections having a loaded leading-edge A4K6 mean line have been investigated systematically in a two-dimensional porous-wall cascade over a range of Reynolds numbers from 160,000 to 385,000. Blades cambered to have isolated-airfoil lift coefficients of 0.6, 1.2, 1.8, and 2.4 were tested over the usable angle-of-attack range at inlet-air angles of 30 degrees, 45 degrees, and 60 degrees and solidities of 1.0 and 1.5. A comparison with data of NACA RM L51G31, shows that the angle-of-attack operating range is 2 degrees to 4 degrees less than the range for the uniformly loaded section; however, the wake losses near design angle of attack are slightly lower than those for the uniformly loaded section. Except for highly cambered blades at high inlet angles, the 63-(C s oA4K6)06 compressor-blade sections are capable of more efficient operation for moderate-speed subsonic compressors at design angle of attack than are the 65-(C s oa10)10 or the 65-(c s oA2I8b)10 compressor-blade sections. In contrast to the other sections, the loaded leading-edge sections are capable of operating efficiently at the lower Reynolds numbers.

Emery, James C↗