Application of Micropore Filter Technology: Exploring the Blood Flow Path in Arterial-Line Filters and Its Effect on Bubble Trapping Functions

PubMed Central

Herbst, Daniel P.

2017-01-01

Abstract: Conventional arterial-line filters commonly use a large volume circular shaped housing, a wetted micropore screen, and a purge port to trap, separate, and remove gas bubbles from extracorporeal blood flow. Focusing on the bubble trapping function, this work attempts to explore how the filter housing shape and its resulting blood flow path affect the clinical application of arterial-line filters in terms of gross air handling. A video camera was used in a wet-lab setting to record observations made during gross air-bolus injections in three different radially designed filters using a 30–70% glycerol–saline mixture flowing at 4.5 L/min. Two of the filters both had inlet ports attached near the filter-housing top with bottom oriented outlet ports at the bottom, whereas the third filter had its inlet and outlet ports both located at the bottom of the filter housing. The two filters with top-in bottom-out fluid paths were shown to direct the incoming flow downward as it passed through the filter, placing the forces of buoyancy and viscous drag in opposition to each other. This contrasted with the third filter's bottom-in bottom-out fluid path, which was shown to direct the incoming flow upward so that the forces of buoyancy and viscous drag work together. The direction of the blood flow path through a filter may be important to the application of arterial-line filter technology as it helps determine how the forces of buoyancy and flow are aligned with one another. PMID:28298665

Application of Micropore Filter Technology: Exploring the Blood Flow Path in Arterial-Line Filters and Its Effect on Bubble Trapping Functions.

PubMed

Herbst, Daniel P

2017-03-01

Conventional arterial-line filters commonly use a large volume circular shaped housing, a wetted micropore screen, and a purge port to trap, separate, and remove gas bubbles from extracorporeal blood flow. Focusing on the bubble trapping function, this work attempts to explore how the filter housing shape and its resulting blood flow path affect the clinical application of arterial-line filters in terms of gross air handling. A video camera was used in a wet-lab setting to record observations made during gross air-bolus injections in three different radially designed filters using a 30-70% glycerol-saline mixture flowing at 4.5 L/min. Two of the filters both had inlet ports attached near the filter-housing top with bottom oriented outlet ports at the bottom, whereas the third filter had its inlet and outlet ports both located at the bottom of the filter housing. The two filters with top-in bottom-out fluid paths were shown to direct the incoming flow downward as it passed through the filter, placing the forces of buoyancy and viscous drag in opposition to each other. This contrasted with the third filter's bottom-in bottom-out fluid path, which was shown to direct the incoming flow upward so that the forces of buoyancy and viscous drag work together. The direction of the blood flow path through a filter may be important to the application of arterial-line filter technology as it helps determine how the forces of buoyancy and flow are aligned with one another.

Shape Memory Actuated Normally Open Permanent Isolation Valve

NASA Technical Reports Server (NTRS)

Ramspacher, Daniel J. (Inventor); Bacha, Caitlin E. (Inventor)

2017-01-01

A valve assembly for an in-space propulsion system includes an inlet tube, an outlet tube, a valve body coupling the inlet tube to the outlet tube and defining a propellant flow path, a valve stem assembly disposed within the valve body, an actuator body coupled to the valve body, the valve stem assembly extending from an interior of the valve body to an interior of the actuator body, and an actuator assembly disposed within the actuator body and coupled to the valve stem assembly, the actuator assembly including a shape memory actuator member that when heated to a transition temperature is configured to enable the valve stem assembly to engage the outlet tube and seal the propellant flow path.

Flow-Cell-Induced Dispersion in Flow-through Absorbance Detection Systems: True Column Effluent Peak Variance.

PubMed

Dasgupta, Purnendu K; Shelor, Charles Phillip; Kadjo, Akinde Florence; Kraiczek, Karsten G

2018-02-06

Following a brief overview of the emergence of absorbance detection in liquid chromatography, we focus on the dispersion caused by the absorbance measurement cell and its inlet. A simple experiment is proposed wherein chromatographic flow and conditions are held constant but a variable portion of the column effluent is directed into the detector. The temporal peak variance (σ t,obs 2 ), which increases as the flow rate (F) through the detector decreases, is found to be well-described as a quadratic function of 1 / F . This allows the extrapolation of the results to zero residence time in the detector and thence the determination of the true variance of the peak prior to the detector (this includes contribution of all preceding components). This general approach should be equally applicable to detection systems other than absorbance. We also experiment where the inlet/outlet system remains the same but the path length is varied. This allows one to assess the individual contributions of the cell itself and the inlet/outlet system.to the total observed peak. The dispersion in the cell itself has often been modeled as a flow-independent parameter, dependent only on the cell volume. Except for very long path/large volume cells, this paradigm is simply incorrect.

Problems in creation of modern air inlet filters of power gas turbine plants in Russia and methods of their solving

NASA Astrophysics Data System (ADS)

Mikhaylov, V. E.; Khomenok, L. A.; Sherapov, V. V.

2016-08-01

The main problems in creation and operation of modern air inlet paths of gas turbine plants installed as part of combined-cycle plants in Russia are presented. It is noted that design features of air inlet filters shall be formed at the stage of the technical assignment not only considering the requirements of gas turbine plant manufacturer but also climatic conditions, local atmospheric air dustiness, and a number of other factors. The recommendations on completing of filtration system for air inlet filter of power gas turbine plants depending on the facility location are given, specific defects in design and experience in operation of imported air inlet paths are analyzed, and influence of cycle air preparation quality for gas turbine plant on value of operating expenses and cost of repair works is noted. Air treatment equipment of various manufacturers, influence of aerodynamic characteristics on operation of air inlet filters, features of filtration system operation, anti-icing system, weather canopies, and other elements of air inlet paths are considered. It is shown that nonuniformity of air flow velocity fields in clean air chamber has a negative effect on capacity and aerodynamic resistance of air inlet filter. Besides, the necessity in installation of a sufficient number of differential pressure transmitters allowing controlling state of each treatment stage not being limited to one measurement of total differential pressure in the filtration system is noted in the article. According to the results of the analysis trends and methods for modernization of available equipment for air inlet path, the importance of creation and implementation of new technologies for manufacturing of filtering elements on sites of Russia within the limits of import substitution are given, and measures on reliability improvement and energy efficiency for air inlet filter are considered.

High performance internal reforming unit for high temperature fuel cells

DOEpatents

Ma, Zhiwen [Sandy Hook, CT; Venkataraman, Ramakrishnan [New Milford, CT; Novacco, Lawrence J [Brookfield, CT

2008-10-07

A fuel reformer having an enclosure with first and second opposing surfaces, a sidewall connecting the first and second opposing surfaces and an inlet port and an outlet port in the sidewall. A plate assembly supporting a catalyst and baffles are also disposed in the enclosure. A main baffle extends into the enclosure from a point of the sidewall between the inlet and outlet ports. The main baffle cooperates with the enclosure and the plate assembly to establish a path for the flow of fuel gas through the reformer from the inlet port to the outlet port. At least a first directing baffle extends in the enclosure from one of the sidewall and the main baffle and cooperates with the plate assembly and the enclosure to alter the gas flow path. Desired graded catalyst loading pattern has been defined for optimized thermal management for the internal reforming high temperature fuel cells so as to achieve high cell performance.

Opposed slant tube diabatic sorber

DOEpatents

Erickson, Donald C.

2004-01-20

A sorber comprised of at least three concentric coils of tubing contained in a shell with a flow path for liquid sorbent in one direction, a flow path for heat transfer fluid which is in counter-current heat exchange relationship with sorbent flow, a sorbate vapor port in communication with at least one of sorbent inlet or exit ports, wherein each coil is coiled in opposite direction to those coils adjoining it, whereby the opposed slant tube configuration is achieved, with structure for flow modification in the core space inside the innermost coil.

Bioreactor for acid mine drainage control

DOEpatents

Zaluski, Marek H.; Manchester, Kenneth R.

2001-01-01

A bioreactor for reacting an aqueous heavy metal and sulfate containing mine drainage solution with sulfate reducing bacteria to produce heavy metal sulfides and reduce the sulfuric acid content of the solution. The reactor is an elongated, horizontal trough defining an inlet section and a reaction section. An inlet manifold adjacent the inlet section distributes aqueous mine drainage solution into the inlet section for flow through the inlet section and reaction section. A sulfate reducing bacteria and bacteria nutrient composition in the inlet section provides sulfate reducing bacteria that with the sulfuric acid and heavy metals in the solution to form solid metal sulfides. The sulfate reducing bacteria and bacteria nutrient composition is retained in the cells of a honeycomb structure formed of cellular honeycomb panels mounted in the reactor inlet section. The honeycomb panels extend upwardly in the inlet section at an acute angle with respect to the horizontal. The cells defined in each panel are thereby offset with respect to the honeycomb cells in each adjacent panel in order to define a tortuous path for the flow of the aqueous solution.

Design modification for the modular helium reactor for higher temperature operation and reliability studies for nuclear hydrogen production processes

NASA Astrophysics Data System (ADS)

Reza, S. M. Mohsin

Design options have been evaluated for the Modular Helium Reactor (MHR) for higher temperature operation. An alternative configuration for the MHR coolant inlet flow path is developed to reduce the peak vessel temperature (PVT). The coolant inlet path is shifted from the annular path between reactor core barrel and vessel wall through the permanent side reflector (PSR). The number and dimensions of coolant holes are varied to optimize the pressure drop, the inlet velocity, and the percentage of graphite removed from the PSR to create this inlet path. With the removal of ˜10% of the graphite from PSR the PVT is reduced from 541°C to 421°C. A new design for the graphite block core has been evaluated and optimized to reduce the inlet coolant temperature with the aim of further reduction of PVT. The dimensions and number of fuel rods and coolant holes, and the triangular pitch have been changed and optimized. Different packing fractions for the new core design have been used to conserve the number of fuel particles. Thermal properties for the fuel elements are calculated and incorporated into these analyses. The inlet temperature, mass flow and bypass flow are optimized to limit the peak fuel temperature (PFT) within an acceptable range. Using both of these modifications together, the PVT is reduced to ˜350°C while keeping the outlet temperature at 950°C and maintaining the PFT within acceptable limits. The vessel and fuel temperatures during low pressure conduction cooldown and high pressure conduction cooldown transients are found to be well below the design limits. The reliability and availability studies for coupled nuclear hydrogen production processes based on the sulfur iodine thermochemical process and high temperature electrolysis process have been accomplished. The fault tree models for both these processes are developed. Using information obtained on system configuration, component failure probability, component repair time and system operating modes and conditions, the system reliability and availability are assessed. Required redundancies are made to improve system reliability and to optimize the plant design for economic performance. The failure rates and outage factors of both processes are found to be well below the maximum acceptable range.

Near-wall serpentine cooled turbine airfoil

DOEpatents

Lee, Ching-Pang

2014-10-28

A serpentine coolant flow path is formed by inner walls in a cavity between pressure and suction side walls of a turbine airfoil, the cavity partitioned by one or more transverse partitions into a plurality of continuous serpentine cooling flow streams each having a respective coolant inlet.

CFD Code Calibration and Inlet-Fairing Effects On a 3D Hypersonic Powered-Simulation Model

NASA Technical Reports Server (NTRS)

Huebner, Lawrence D.; Tatum, Kenneth E.

1993-01-01

A three-dimensional (3D) computational study has been performed addressing issues related to the wind tunnel testing of a hypersonic powered-simulation model. The study consisted of three objectives. The first objective was to calibrate a state-of-the-art computational fluid dynamics (CFD) code in its ability to predict hypersonic powered-simulation flows by comparing CFD solutions with experimental surface pressure data. Aftbody lower surface pressures were well predicted, but lower surface wing pressures were less accurately predicted. The second objective was to determine the 3D effects on the aftbody created by fairing over the inlet; this was accomplished by comparing the CFD solutions of two closed-inlet powered configurations with a flowing- inlet powered configuration. Although results at four freestream Mach numbers indicate that the exhaust plume tends to isolate the aftbody surface from most forebody flow- field differences, a smooth inlet fairing provides the least aftbody force and moment variation compared to a flowing inlet. The final objective was to predict and understand the 3D characteristics of exhaust plume development at selected points on a representative flight path. Results showed a dramatic effect of plume expansion onto the wings as the freestream Mach number and corresponding nozzle pressure ratio are increased.

Modular assembly for supporting, straining, and directing flow to a core in a nuclear reactor

DOEpatents

Pennell, William E.

1977-01-01

A reactor core support arrangement for supporting, straining, and providing fluid flow to the core and periphery of a nuclear reactor during normal operation. A plurality of removable inlet modular units are contained within permanent liners in the lower supporting plate of the reactor vessel lower internals. During normal operation (1) each inlet modular unit directs main coolant flow to a plurality of core assemblies, the latter being removably supported in receptacles in the upper portion of the modular unit and (2) each inlet modular unit may direct bypass flow to a low pressure annular region of the reactor vessel. Each inlet modular unit may include special fluid seals interposed between mating surfaces of the inlet modular units and the core assemblies and between the inlet modular units and the liners, to minimize leakage and achieve an hydraulic balance. Utilizing the hydraulic balance, the modular units are held in the liners and the assemblies are held in the modular unit receptacles by their own respective weight. Included as part of the permanent liners below the horizontal support plate are generally hexagonal axial debris barriers. The axial debris barriers collectively form a bottom boundary of a secondary high pressure plenum, the upper boundary of which is the bottom surface of the horizontal support plate. Peripheral liners include radial debris barriers which collectively form a barrier against debris entry radially. During normal operation primary coolant inlet openings in the liner, below the axial debris barriers, pass a large amount of coolant into the inlet modular units, and secondary coolant inlet openings in the portion of the liners within the secondary plenum pass a small amount of coolant into the inlet modular units. The secondary coolant inlet openings also provide alternative coolant inlet flow paths in the unlikely event of blockage of the primary inlet openings. The primary inlet openings have characteristics which limit the entry of debris and minimize the potential for debris entering the primary inlets blocking the secondary inlets from inside the modular unit.

Magnetic roller gas gate employing transonic sweep gas flow to isolate regions of differing gaseous composition or pressure

DOEpatents

Doehler, Joachim

1994-12-20

Disclosed herein is an improved gas gate for interconnecting regions of differing gaseous composition and/or pressure. The gas gate includes a narrow, elongated passageway through which substrate material is adapted to move between said regions and inlet means for introducing a flow of non-contaminating sweep gas into a central portion of said passageway. The gas gate is characterized in that the height of the passageway and the flow rate of the sweep gas therethrough provides for transonic flow of the sweep gas between the inlet means and at least one of the two interconnected regions, thereby effectively isolating one region, characterized by one composition and pressure, from another region, having a differing composition and/or pressure, by decreasing the mean-free-path length between collisions of diffusing species within the transonic flow region. The gas gate preferably includes a manifold at the juncture point where the gas inlet means and the passageway interconnect.

Jet Engine Fan Response to Inlet Distortions Generated by Ingesting Boundary Layer Flow

NASA Astrophysics Data System (ADS)

Giuliani, James Edward

Future civil transport designs may incorporate engines integrated into the body of the aircraft to take advantage of efficiency increases due to weight and drag reduction. Additional increases in engine efficiency are predicted if the inlets ingest the lower momentum boundary layer flow that develops along the surface of the aircraft. Previous studies have shown, however, that the efficiency benefits of Boundary Layer Ingesting (BLI) inlets are very sensitive to the magnitude of fan and duct losses, and blade structural response to the non-uniform flow field that results from a BLI inlet has not been studied in-depth. This project represents an effort to extend the modeling capabilities of TURBO, an existing rotating turbomachinery unsteady analysis code, to include the ability to solve the external and internal flow fields of a BLI inlet. The TURBO code has been a successful tool in evaluating fan response to flow distortions for traditional engine/inlet integrations. Extending TURBO to simulate the external and inlet flow field upstream of the fan will allow accurate pressure distortions that result from BLI inlet configurations to be computed and used to analyze fan aerodynamics and structural response. To validate the modifications for the BLI inlet flow field, an experimental NASA project to study flush-mounted S-duct inlets with large amounts of boundary layer ingestion was modeled. Results for the flow upstream and in the inlet are presented and compared to experimental data for several high Reynolds number flows to validate the modifications to the solver. Once the inlet modifications were validated, a hypothetical compressor fan was connected to the inlet, matching the inlet operating conditions so that the effect on the distortion could be evaluated. Although the total pressure distortion upstream of the fan was symmetrical for this geometry, the pressure rise generated by the fan blades was not, because of the velocity non-uniformity of the distortion. Total pressure profiles at various axial locations are computed to identify the overall distortion pattern, how the distortion evolves through the blade passages and mixes out downstream of the blades, and where any critical performance concerns might be. Stall cells are identified that are stationary in the absolute frame and are fixed to the inlet distortion. Flow paths around the blades are examined to study the stall mechanism. Rather than a static airfoil stall, it is observed that the non-uniform pressure loading promotes a three-dimensional dynamic stall. The stall occurs at a point of rapid incidence angle oscillation, observed when a blade passes through the distortion, and re-attaches when the blade leaves the distortion.

CFD code calibration and inlet-fairing effects on a 3D hypersonic powered-simulation model

NASA Technical Reports Server (NTRS)

Huebner, Lawrence D.; Tatum, Kenneth E.

1993-01-01

A three-dimensional (3D) computational study has been performed addressing issues related to the wind tunnel testing of a hypersonic powered-simulation model. The study consisted of three objectives. The first objective was to calibrate a state-of-the-art computational fluid dynamics (CFD) code in its ability to predict hypersonic powered-simulation flows by comparing CFD solutions with experimental surface pressure dam. Aftbody lower surface pressures were well predicted, but lower surface wing pressures were less accurately predicted. The second objective was to determine the 3D effects on the aftbody created by fairing over the inlet; this was accomplished by comparing the CFD solutions of two closed-inlet powered configurations with a flowing-inlet powered configuration. Although results at four freestream Mach numbers indicate that the exhaust plume tends to isolate the aftbody surface from most forebody flowfield differences, a smooth inlet fairing provides the least aftbody force and moment variation compared to a flowing inlet. The final objective was to predict and understand the 3D characteristics of exhaust plume development at selected points on a representative flight path. Results showed a dramatic effect of plume expansion onto the wings as the freestream Mach number and corresponding nozzle pressure ratio are increased.

Quasi-2D Unsteady Flow Procedure for Real Fluids (PREPRINT)

DTIC Science & Technology

2006-05-17

water /steam/ oil piping networks, refinery systems, gas-turbine secondary flow -path and cooling networks...friction factor, f, which is a function of the local Reynolds number and the wall surface roughness . For the viscous flow examples presented below, the...3.5 4 4.5 Time ( s ) V el oc ity (m / s ) Line 2 Inlet 25% 50% 75% Exit Velocity Figure 4. Water transient viscous pipe flow using

Effects of Temperature on the Performance of a Small Internal Combustion Engine at Altitude

DTIC Science & Technology

2013-03-21

flexible diaphragm was attached to damp out pulses in the air flow pulsations . Their method of temperature control was electric heating of the intake air...42  Figure 14. Heat exchanger ................................................................................................ 45  Figure...15. Both liquid nitrogen lines from Dewar ............................................................ 45  Figure 16. Engine inlet flow path heat

Diesel engine exhaust oxidizer

DOE Office of Scientific and Technical Information (OSTI.GOV)

Kammel, R.A.

1992-06-16

This patent describes a diesel engine exhaust oxidizing device. It comprises: an enclosure having an inlet for receiving diesel engine exhaust, a main flow path through the enclosure to an outlet of the enclosure, a by-ass through the enclosure, and a microprocessor control means.

Technology Horizons: A Vision for Air Force Science and Technology 2010-30

DTIC Science & Technology

2011-09-01

software, hardware, and networks, it is now recognized as en- compassing the entire system that couples information flow and decision processes across...acceleration, and scramjet cruise. Inward turning inlets and a dual- flow path design allow high volumetric efficiency, and high cruise speed provides...the same time, emerging “third- stream engine architectures” can enable constant-mass- flow engines that can provide further reductions in fuel

Integrated packaging of multiple double sided cooling planar bond power modules

DOE Office of Scientific and Technical Information (OSTI.GOV)

Liang, Zhenxian

An integrated double sided cooled power module has one or multiple phase legs configuration including one or more planar power packages, each planar power package having an upper power switch unit and a lower power switch unit directly bonded and interconnected between two insulated power substrates, and further sandwiched between two heat exchangers via direct bonds. A segmented coolant manifold is interposed with the one or more planar power packages and creates a sealed enclosure that defines a coolant inlet, a coolant outlet and a coolant flow path between the inlet and the outlet. A coolant circulates along the flowmore » path to remove heat and increase the power density of the power module.« less

Computational fluid dynamics analysis of SSME phase 2 and phase 2+ preburner injector element hydrogen flow paths

NASA Technical Reports Server (NTRS)

Ruf, Joseph H.

1992-01-01

Phase 2+ Space Shuttle Main Engine powerheads, E0209 and E0215 degraded their main combustion chamber (MCC) liners at a faster rate than is normal for phase 2 powerheads. One possible cause of the accelerated degradation was a reduction of coolant flow through the MCC. Hardware changes were made to the preburner fuel leg which may have reduced the resistance and, therefore, pulled some of the hydrogen from the MCC coolant leg. A computational fluid dynamics (CFD) analysis was performed to determine hydrogen flow path resistances of the phase 2+ fuel preburner injector elements relative to the phase 2 element. FDNS was implemented on axisymmetric grids with the hydrogen assumed to be incompressible. The analysis was performed in two steps: the first isolated the effect of the different inlet areas and the second modeled the entire injector element hydrogen flow path.

A water tunnel flow visualization study of the F-15

NASA Technical Reports Server (NTRS)

Lorincz, D. J.

1978-01-01

Water tunnel studies were performed to qualitatively define the flow field of the F-15 aircraft. Two lengthened forebodies, one with a modified cross-sectional shape, were tested in addition to the basic forebody. Particular emphasis was placed on defining vortex flows generated at high angles of attack. The flow visualization tests were conducted in the Northrop diagnostic water tunnel using a 1/48-scale model of the F-15. Flow visualization pictures were obtained over an angle-of-attack range to 55 deg and sideslip angles up to 10 deg. The basic aircraft configuration was investigated in detail to determine the vortex flow field development, vortex path, and vortex breakdown characteristics as a function of angle of attack and sideslip. Additional tests showed that the wing upper surface vortex flow fields were sensitive to variations in inlet mass flow ratio and inlet cowl deflection angle. Asymmetries in the vortex systems generated by each of the three forebodies were observed in the water tunnel at zero sideslip and high angles of attack.

Assembly for directing combustion gas

DOE Office of Scientific and Technical Information (OSTI.GOV)

Charron, Richard C.; Little, David A.; Snyder, Gary D.

2016-04-12

An arrangement is provided for delivering gases from a plurality of combustors of a can-annular gas turbine combustion engine to a first row of turbine blades including a first row of turbine blades. The arrangement includes a gas path cylinder, a cone and an integrated exit piece (IEP) for each combustor. Each IEP comprises an inlet chamber for receiving a gas flow from a respective combustor, and includes a connection segment. The IEPs are connected together to define an annular chamber extending circumferentially and concentric to an engine longitudinal axis, for delivering the gas flow to the first row ofmore » blades. A radiused joint extends radially inward from a radially outer side of the inlet chamber to an outer boundary of the annular chamber, and a flared fillet extends radially inward from a radially inner side of the inlet chamber to an inner boundary of the annular chamber.« less

Linear induction pump

DOEpatents

Meisner, John W.; Moore, Robert M.; Bienvenue, Louis L.

1985-03-19

Electromagnetic linear induction pump for liquid metal which includes a unitary pump duct. The duct comprises two substantially flat parallel spaced-apart wall members, one being located above the other and two parallel opposing side members interconnecting the wall members. Located within the duct are a plurality of web members interconnecting the wall members and extending parallel to the side members whereby the wall members, side members and web members define a plurality of fluid passageways, each of the fluid passageways having substantially the same cross-sectional flow area. Attached to an outer surface of each side member is an electrically conductive end bar for the passage of an induced current therethrough. A multi-phase, electrical stator is located adjacent each of the wall members. The duct, stators, and end bars are enclosed in a housing which is provided with an inlet and outlet in fluid communication with opposite ends of the fluid passageways in the pump duct. In accordance with a preferred embodiment, the inlet and outlet includes a transition means which provides for a transition from a round cross-sectional flow path to a substantially rectangular cross-sectional flow path defined by the pump duct.

The design of an air-cooled metallic high temperature radial turbine

NASA Technical Reports Server (NTRS)

Snyder, Philip H.; Roelke, Richard J.

1988-01-01

Recent trends in small advanced gas turbine engines call for higher turbine inlet temperatures. Advances in radial turbine technology have opened the way for a cooled metallic radial turbine capable of withstanding turbine inlet temperatures of 2500 F while meeting the challenge of high efficiency in this small flow size range. In response to this need, a small air-cooled radial turbine has been designed utilizing internal blade coolant passages. The coolant flow passage design is uniquely tailored to simultaneously meet rotor cooling needs and rotor fabrication constraints. The rotor flow-path design seeks to realize improved aerodynamic blade loading characteristics and high efficiency while satisfying rotor life requirements. An up-scaled version of the final engine rotor is currently under fabrication and, after instrumentation, will be tested in the warm turbine test facility at the NASA Lewis Research Center.

HOST turbine heat transfer subproject overview

NASA Technical Reports Server (NTRS)

Gladden, Herbert J.

1986-01-01

The experimental part of the turbine heat transfer subproject consists of six large experiments, which are highlighted in this overview, and three of somewhat more modest scope. One of the initial efforts was the stator airfoil heat transfer program. The non-film cooled and the showerhead film cooled data have already been reported. The gill region film cooling effort is currently underway. The investigation of secondary flows in a 90 deg curved duct, was completed. The first phase examined flows with a relatively thin inlet boundary layer and low free stream turbulence. The second phase studied a thicker inlet boundary layer and higher free stream turbulence. A comparison of analytical and experimental cross flow velocity vectors is shown for the 60 deg plane. Two experiments were also conducted in the high pressure facility. One examined full coverage film cooled vanes, and the other, advanced instrumentation. The other three large experimental efforts were conducted in a rotation reference frame. An experiment to obtain gas path airfoil heat transfer coefficients in the large, low speed turbine was completed. Single-stage data with both high and low-inlet turbulence were taken. The second phase examined a one and one-half stage turbine and focused on the second vane row. Under phase 3 aerodynamic quantities such as interrow time-averaged and rms values of velocity, flow angle, inlet turbulence, and surface pressure distribution were measured.

Heat exchanger with ceramic elements

DOEpatents

Corey, John A.

1986-01-01

An annular heat exchanger assembly includes a plurality of low thermal growth ceramic heat exchange members with inlet and exit flow ports on distinct faces. A mounting member locates each ceramic member in a near-annular array and seals the flow ports on the distinct faces into the separate flow paths of the heat exchanger. The mounting member adjusts for the temperature gradient in the assembly and the different coefficients of thermal expansion of the members of the assembly during all operating temperatures.

Dissolution-induced preferential flow in a limestone fracture.

PubMed

Liu, Jishan; Polak, Amir; Elsworth, Derek; Grader, Avrami

2005-06-01

Flow in a rock fracture is surprisingly sensitive to the evolution of flow paths that develop as a result of dissolution. Net dissolution may either increase or decrease permeability uniformly within the fracture, or may form a preferential flow path through which most of the injected fluid flows, depending on the prevailing ambient mechanical and chemical conditions. A flow-through test was completed on an artificial fracture in limestone at room temperature under ambient confining stress of 3.5 MPa. The sample was sequentially circulated by water of two different compositions through the 1500 h duration of the experiment; the first 935 h by tap groundwater, followed by 555 h of distilled water. Measurements of differential pressures between the inlet and the outlet, fluid and dissolved mass fluxes, and concurrent X-ray CT imaging and sectioning were used to characterize the evolution of flow paths within the limestone fracture. During the initial circulation of groundwater, the differential pressure increased almost threefold, and was interpreted as a net reduction in permeability as the contacting asperities across the fracture are removed, and the fracture closes. With the circulation of distilled water, permeability initially reduces threefold, and ultimately increases by two orders of magnitude. This spontaneous switch from net decrease in permeability, to net increase occurred with no change in flow rate or applied effective stress, and is attributed to the evolving localization of flow path as evidenced by CT images. Based on the X-ray CT characterizations, a flow path-dependent flow model was developed to simulate the evolution of flow paths within the fracture and its influence on the overall flow behaviors of the injected fluid in the fracture.

Serpentine Diffuser Performance with Emphasis on Future Introduction to a Transonic Fan (Postprint)

DTIC Science & Technology

2013-01-01

conditioning barrel . The velocity distribution across the flow conditioning barrel was measured at the same axial location of inlet temperature and...rakes at the same axial plane (AIP) of the total pressure probe tips. The probes were constructed from stainless steel tubing with 0.027 inch inside...numbers with 195 axial and circumferential static pressure measurements within the diffuser flow path. Pressure distortion at the diffuser discharge

CFD validation experiments for hypersonic flows

NASA Technical Reports Server (NTRS)

Marvin, Joseph G.

1992-01-01

A roadmap for CFD code validation is introduced. The elements of the roadmap are consistent with air-breathing vehicle design requirements and related to the important flow path components: forebody, inlet, combustor, and nozzle. Building block and benchmark validation experiments are identified along with their test conditions and measurements. Based on an evaluation criteria, recommendations for an initial CFD validation data base are given and gaps identified where future experiments could provide new validation data.

Numerical Investigation of Different Radial Inlet Forms for Centrifugal Compressor and Influence of the Deflectors Number by Means of Computational Fluid Dynamics Methods with Computational Model Validation

NASA Astrophysics Data System (ADS)

Kozhukhov, Y. V.; Yun, V. K.; Reshetnikova, L. V.; Prokopovich, M. V.

2015-08-01

The goal of this work is numerical experiments for five different types of the centrifugal compressor's inlet chambers with the help of CFD-methods and comparison of the computational results with the results of the real experiment which was held in the Nevskiy Lenin Plant in Saint-Petersburg. In the context of one of the chambers the influence of deflectors on its characteristics was investigated. The objects of investigation are 5 inlet chambers of different types which differ from each other by deflectors’ existence and by its number. The comparative analyze of the results of numerical and real experiments was held by means of comparison of relative velocity and static pressure coefficient distribution on hub and shroud region, and also by means of loss coefficient values change for all five chambers. As a result of the numerical calculation the quantitative and qualitative departure of CFD- calculations results and real experiment were found out. The investigation of the influence of the number of deflectors on flow parameters was carried out. The results of the study prove that the presence of the deflectors on flow path significantly increases the probability of the flow separations and reversed flows appearance on them. At the same time, the complete absence of the deflectors in the chamber significantly increases circumferential distortion of the flow; however the loss coefficient decreases anyway, the high values of which are caused by the shock flow existence. Thus, the profiling of the deflectors of the inlet chamber should be given a special attention.

Shock Position Control for Mode Transition in a Turbine Based Combined Cycle Engine Inlet Model

NASA Technical Reports Server (NTRS)

Csank, Jeffrey T.; Stueber, Thomas J.

2013-01-01

A dual flow-path inlet for a turbine based combined cycle (TBCC) propulsion system is to be tested in order to evaluate methodologies for performing a controlled inlet mode transition. Prior to experimental testing, simulation models are used to test, debug, and validate potential control algorithms which are designed to maintain shock position during inlet disturbances. One simulation package being used for testing is the High Mach Transient Engine Cycle Code simulation, known as HiTECC. This paper discusses the development of a mode transition schedule for the HiTECC simulation that is analogous to the development of inlet performance maps. Inlet performance maps, derived through experimental means, describe the performance and operability of the inlet as the splitter closes, switching power production from the turbine engine to the Dual Mode Scram Jet. With knowledge of the operability and performance tradeoffs, a closed loop system can be designed to optimize the performance of the inlet. This paper demonstrates the design of the closed loop control system and benefit with the implementation of a Proportional-Integral controller, an H-Infinity based controller, and a disturbance observer based controller; all of which avoid inlet unstart during a mode transition with a simulated disturbance that would lead to inlet unstart without closed loop control.

Advanced Shock Position Control for Mode Transition in a Turbine Based Combined Cycle Engine Inlet Model

NASA Technical Reports Server (NTRS)

Csank, Jeffrey T.; Stueber, Thomas J.

2013-01-01

A dual flow-path inlet system is being tested to evaluate methodologies for a Turbine Based Combined Cycle (TBCC) propulsion system to perform a controlled inlet mode transition. Prior to experimental testing, simulation models are used to test, debug, and validate potential control algorithms. One simulation package being used for testing is the High Mach Transient Engine Cycle Code simulation, known as HiTECC. This paper discusses the closed loop control system, which utilizes a shock location sensor to improve inlet performance and operability. Even though the shock location feedback has a coarse resolution, the feedback allows for a reduction in steady state error and, in some cases, better performance than with previous proposed pressure ratio based methods. This paper demonstrates the design and benefit with the implementation of a proportional-integral controller, an H-Infinity based controller, and a disturbance observer based controller.

Combustor assembly in a gas turbine engine

DOEpatents

Wiebe, David J; Fox, Timothy A

2013-02-19

A combustor assembly in a gas turbine engine. The combustor assembly includes a combustor device coupled to a main engine casing, a first fuel injection system, a transition duct, and an intermediate duct. The combustor device includes a flow sleeve for receiving pressurized air and a liner disposed radially inwardly from the flow sleeve. The first fuel injection system provides fuel that is ignited with the pressurized air creating first working gases. The intermediate duct is disposed between the liner and the transition duct and defines a path for the first working gases to flow from the liner to the transition duct. An intermediate duct inlet portion is associated with a liner outlet and allows movement between the intermediate duct and the liner. An intermediate duct outlet portion is associated with a transition duct inlet section and allows movement between the intermediate duct and the transition duct.

A CFD validation roadmap for hypersonic flows

NASA Technical Reports Server (NTRS)

Marvin, Joseph G.

1992-01-01

A roadmap for computational fluid dynamics (CFD) code validation is developed. The elements of the roadmap are consistent with air-breathing vehicle design requirements and related to the important flow path components: forebody, inlet, combustor, and nozzle. Building block and benchmark validation experiments are identified along with their test conditions and measurements. Based on an evaluation criteria, recommendations for an initial CFD validation data base are given and gaps identified where future experiments would provide the needed validation data.

A CFD validation roadmap for hypersonic flows

NASA Technical Reports Server (NTRS)

Marvin, Joseph G.

1993-01-01

A roadmap for computational fluid dynamics (CFD) code validation is developed. The elements of the roadmap are consistent with air-breathing vehicle design requirements and related to the important flow path components: forebody, inlet, combustor, and nozzle. Building block and benchmark validation experiments are identified along with their test conditions and measurements. Based on an evaluation criteria, recommendations for an initial CFD validation data base are given and gaps identified where future experiments would provide the needed validation data.

Carbon nanotube heat-exchange systems

DOEpatents

Hendricks, Terry Joseph; Heben, Michael J.

2008-11-11

A carbon nanotube heat-exchange system (10) and method for producing the same. One embodiment of the carbon nanotube heat-exchange system (10) comprises a microchannel structure (24) having an inlet end (30) and an outlet end (32), the inlet end (30) providing a cooling fluid into the microchannel structure (24) and the outlet end (32) discharging the cooling fluid from the microchannel structure (24). At least one flow path (28) is defined in the microchannel structure (24), fluidically connecting the inlet end (30) to the outlet end (32) of the microchannel structure (24). A carbon nanotube structure (26) is provided in thermal contact with the microchannel structure (24), the carbon nanotube structure (26) receiving heat from the cooling fluid in the microchannel structure (24) and dissipating the heat into an external medium (19).

Apparatus for sampling and characterizing aerosols

DOEpatents

Dunn, Patrick F.; Herceg, Joseph E.; Klocksieben, Robert H.

1986-01-01

Apparatus for sampling and characterizing aerosols having a wide particle size range at relatively low velocities may comprise a chamber having an inlet and an outlet, the chamber including: a plurality of vertically stacked, successive particle collection stages; each collection stage includes a separator plate and a channel guide mounted transverse to the separator plate, defining a labyrinthine flow path across the collection stage. An opening in each separator plate provides a path for the aerosols from one collection stage to the next. Mounted within each collection stage are one or more particle collection frames.

Flow of GE90 Turbofan Engine Simulated

NASA Technical Reports Server (NTRS)

Veres, Joseph P.

1999-01-01

The objective of this task was to create and validate a three-dimensional model of the GE90 turbofan engine (General Electric) using the APNASA (average passage) flow code. This was a joint effort between GE Aircraft Engines and the NASA Lewis Research Center. The goal was to perform an aerodynamic analysis of the engine primary flow path, in under 24 hours of CPU time, on a parallel distributed workstation system. Enhancements were made to the APNASA Navier-Stokes code to make it faster and more robust and to allow for the analysis of more arbitrary geometry. The resulting simulation exploited the use of parallel computations by using two levels of parallelism, with extremely high efficiency.The primary flow path of the GE90 turbofan consists of a nacelle and inlet, 49 blade rows of turbomachinery, and an exhaust nozzle. Secondary flows entering and exiting the primary flow path-such as bleed, purge, and cooling flows-were modeled macroscopically as source terms to accurately simulate the engine. The information on these source terms came from detailed descriptions of the cooling flow and from thermodynamic cycle system simulations. These provided boundary condition data to the three-dimensional analysis. A simplified combustor was used to feed boundary conditions to the turbomachinery. Flow simulations of the fan, high-pressure compressor, and high- and low-pressure turbines were completed with the APNASA code.

Process optimization of helium cryo plant operation for SST-1 superconducting magnet system

NASA Astrophysics Data System (ADS)

Panchal, P.; Panchal, R.; Patel, R.; Mahesuriya, G.; Sonara, D.; Srikanth G, L. N.; Garg, A.; Christian, D.; Bairagi, N.; Sharma, R.; Patel, K.; Shah, P.; Nimavat, H.; Purwar, G.; Patel, J.; Tanna, V.; Pradhan, S.

2017-02-01

Several plasma discharge campaigns have been carried out in steady state superconducting tokamak (SST-1). SST-1 has toroidal field (TF) and poloidal field (PF) superconducting magnet system (SCMS). The TF coils system is cooled to 4.5 - 4.8 K at 1.5 - 1.7 bar(a) under two phase flow condition using 1.3 kW helium cryo plant. Experience revealed that the PF coils demand higher pressure heads even at lower temperatures in comparison to TF coils because of its longer hydraulic path lengths. Thermal run away are observed within PF coils because of single common control valve for all PF coils in distribution system having non-uniform lengths. Thus it is routine practice to stop the cooling of PF path and continue only TF cooling at SCMS inlet temperature of ˜ 14 K. In order to achieve uniform cool down, different control logic is adopted to make cryo stable system. In adopted control logic, the SCMS are cooled down to 80 K at constant inlet pressure of 9 bar(a). After authorization of turbine A/B, the SCMS inlet pressure is gradually controlled by refrigeration J-T valve to achieve stable operation window for cryo system. This paper presents process optimization for cryo plant operation for SST-1 SCMS.

Computational Fluid Dynamics Analysis Method Developed for Rocket-Based Combined Cycle Engine Inlet

NASA Technical Reports Server (NTRS)

1997-01-01

Renewed interest in hypersonic propulsion systems has led to research programs investigating combined cycle engines that are designed to operate efficiently across the flight regime. The Rocket-Based Combined Cycle Engine is a propulsion system under development at the NASA Lewis Research Center. This engine integrates a high specific impulse, low thrust-to-weight, airbreathing engine with a low-impulse, high thrust-to-weight rocket. From takeoff to Mach 2.5, the engine operates as an air-augmented rocket. At Mach 2.5, the engine becomes a dual-mode ramjet; and beyond Mach 8, the rocket is turned back on. One Rocket-Based Combined Cycle Engine variation known as the "Strut-Jet" concept is being investigated jointly by NASA Lewis, the U.S. Air Force, Gencorp Aerojet, General Applied Science Labs (GASL), and Lockheed Martin Corporation. Work thus far has included wind tunnel experiments and computational fluid dynamics (CFD) investigations with the NPARC code. The CFD method was initiated by modeling the geometry of the Strut-Jet with the GRIDGEN structured grid generator. Grids representing a subscale inlet model and the full-scale demonstrator geometry were constructed. These grids modeled one-half of the symmetric inlet flow path, including the precompression plate, diverter, center duct, side duct, and combustor. After the grid generation, full Navier-Stokes flow simulations were conducted with the NPARC Navier-Stokes code. The Chien low-Reynolds-number k-e turbulence model was employed to simulate the high-speed turbulent flow. Finally, the CFD solutions were postprocessed with a Fortran code. This code provided wall static pressure distributions, pitot pressure distributions, mass flow rates, and internal drag. These results were compared with experimental data from a subscale inlet test for code validation; then they were used to help evaluate the demonstrator engine net thrust.

The effect of geometry and operation conditions on the performance of a gas-liquid cylindrical cyclone separator with new structure

NASA Astrophysics Data System (ADS)

Han, Qing; Zhang, Chi; Xu, Bo; Chen, Jiangping

2013-07-01

The hydrodynamic flow behavior, effects of geometry and working conditions of a gas-liquid cylindrical cyclone separator with a new structure are investigated by computational fluid dynamic and experiment. Gas liquid cylindrical cyclone separator is widely used in oil industry, refrigeration system because of its simple structure, high separating efficiency, little maintenance and no moving parts nor internal devices. In this work, a gas liquid cylindrical cyclone separator with new structure used before evaporator in refrigeration system can remove the vapor from the mixture and make evaporator compact by improving its heat exchange efficiency with the lower inlet quality. It also decreases evaporator pressure drop and reduces compressor work. The two pipes are placed symmetrically which makes each of them can be treated as inlet. It means when the fluids flow reverse, the separator performance will not be influence. Four samples with different geometry parameters are tested by experiment with different inlet quality (0.18-0.33), inlet mass flow rate (65-100kg/h). Compared with the experimental data, CFD simulation results show a good agreement. Eulerian multiphase model and Reynolds Stress Turbulence model are applied in the CFD simulation and obtained the inner flow field such as phase path lines, tangential velocity profiles and pressure and volume of fraction distribution contours. The separator body diameter (24, 36, 48mm) and inlet diameter (3.84, 4.8, 5.76mm) decide the maximum tangential velocity which results in the centrifugal force. The tangential velocity profiles are simulated and compared among different models. The higher tangential velocity makes higher quality of gas outlet but high pressure drop at the same time. Decreasing the inlet diameter increases quality of gas outlet pipe and pressure drop. High gas outlet quality is cost at high pressure drop. Increasing of separator diameter makes gas outlet quality increase first and then decrease but the pressure drop decreases all the way. The offset (0, 2.4, 3.6mm) of gas outlet is an insensitive factor which influences the quality and pressure drop little.

Fuel cell with metal screen flow-field

DOEpatents

Wilson, M.S.; Zawodzinski, C.

1998-08-25

A polymer electrolyte membrane (PEM) fuel cell is provided with electrodes supplied with a reactant on each side of a catalyzed membrane assembly (CMA). The fuel cell includes a metal mesh defining a rectangular flow-field pattern having an inlet at a first corner and an outlet at a second corner located on a diagonal from the first corner, wherein all flow paths from the inlet to the outlet through the square flow field pattern are equivalent to uniformly distribute the reactant over the CMA. In a preferred form of metal mesh, a square weave screen forms the flow-field pattern. In a particular characterization of the present invention, a bipolar plate electrically connects adjacent fuel cells, where the bipolar plate includes a thin metal foil having an anode side and a cathode side; a first metal mesh on the anode side of the thin metal foil; and a second metal mesh on the cathode side of the thin metal foil. In another characterization of the present invention, a cooling plate assembly cools adjacent fuel cells, where the cooling plate assembly includes an anode electrode and a cathode electrode formed of thin conducting foils; and a metal mesh flow field there between for distributing cooling water flow over the electrodes to remove heat generated by the fuel cells. 11 figs.

Fuel cell with metal screen flow-field

DOEpatents

Wilson, Mahlon S.; Zawodzinski, Christine

2001-01-01

A polymer electrolyte membrane (PEM) fuel cell is provided with electrodes supplied with a reactant on each side of a catalyzed membrane assembly (CMA). The fuel cell includes a metal mesh defining a rectangular flow-field pattern having an inlet at a first corner and an outlet at a second corner located on a diagonal from the first corner, wherein all flow paths from the inlet to the outlet through the square flow field pattern are equivalent to uniformly distribute the reactant over the CMA. In a preferred form of metal mesh, a square weave screen forms the flow-field pattern. In a particular characterization of the present invention, a bipolar plate electrically connects adjacent fuel cells, where the bipolar plate includes a thin metal foil having an anode side and a cathode side; a first metal mesh on the anode side of the thin metal foil; and a second metal mesh on the cathode side of the thin metal foil. In another characterization of the present invention, a cooling plate assembly cools adjacent fuel cells, where the cooling plate assembly includes an anode electrode and a cathode electrode formed of thin conducting foils; and a metal mesh flow field therebetween for distributing cooling water flow over the electrodes to remove heat generated by the fuel cells.

Fuel cell with metal screen flow-field

DOEpatents

Wilson, Mahlon S.; Zawodzinski, Christine

1998-01-01

A polymer electrolyte membrane (PEM) fuel cell is provided with electrodes supplied with a reactant on each side of a catalyzed membrane assembly (CMA). The fuel cell includes a metal mesh defining a rectangular flow-field pattern having an inlet at a first corner and an outlet at a second corner located on a diagonal from the first corner, wherein all flow paths from the inlet to the outlet through the square flow field pattern are equivalent to uniformly distribute the reactant over the CMA. In a preferred form of metal mesh, a square weave screen forms the flow-field pattern. In a particular characterization of the present invention, a bipolar plate electrically connects adjacent fuel cells, where the bipolar plate includes a thin metal foil having an anode side and a cathode side; a first metal mesh on the anode side of the thin metal foil; and a second metal mesh on the cathode side of the thin metal foil. In another characterization of the present invention, a cooling plate assembly cools adjacent fuel cells, where the cooling plate assembly includes an anode electrode and a cathode electrode formed of thin conducting foils; and a metal mesh flow field therebetween for distributing cooling water flow over the electrodes to remove heat generated by the fuel cells.

Flow rate logging seepage meter

NASA Technical Reports Server (NTRS)

Reay, William G. (Inventor); Walthall, Harry G. (Inventor)

1996-01-01

An apparatus for remotely measuring and logging the flow rate of groundwater seepage into surface water bodies. As groundwater seeps into a cavity created by a bottomless housing, it displaces water through an inlet and into a waterproof sealed upper compartment, at which point, the water is collected by a collection bag, which is contained in a bag chamber. A magnet on the collection bag approaches a proximity switch as the collection bag fills, and eventually enables the proximity switch to activate a control circuit. The control circuit then rotates a three-way valve from the collection path to a discharge path, enables a data logger to record the time, and enables a pump, which discharges the water from the collection bag, through the three-way valve and pump, and into the sea. As the collection bag empties, the magnet leaves the proximity of the proximity switch, and the control circuit turns off the pump, resets the valve to provide a collection path, and restarts the collection cycle.

Application of nanoscale zero-valent iron tracer to delineate groundwater flow paths between a screened well and an open well in fractured rock

NASA Astrophysics Data System (ADS)

Chuang, P. Y.; Chiu, Y.; Liou, Y. H.; Teng, M. H.; Chia, Y.

2016-12-01

Fracture flow is of importance for water resources as well as the investigation of contaminant pathways. In this study, a novel characterization approach of nanoscale zero-valent iron (nZVI) tracer test was developed to accurately identify the connecting fracture zones of preferential flow between a screened well and an open well. Iron nanoparticles are magnetic and can be attracted by a magnet. This feature make it possible to design a magnet array for attracting nZVI particles at the tracer inlet to characterize the location of incoming tracer in the observation well. This novel approach was tested at two experiment wells with well hydraulic connectivity in a hydrogeological research station in central Taiwan. A heat-pulse flowmeter can be used to detect changes in flow velocity for delineating permeable fracture zones in the borehole and providing the design basis for the tracer test. Then, the most permeable zone in the injection well was hydraulically isolated by well screen to prevent the injected nZVI particles from being stagnated at the hole bottom. Afterwards, another hydraulic test was implemented to re-examine the hydraulic connectivity between the two wells. When nZVI slurry was injected in the injection well, they migrated through connected permeable fractures to the observation well. A breakthrough curve, observed by the fluid conductivity sensor in the observation well, indicated the arrival of nZVI slurry. The iron nanoparticles attracted to the magnets in the observation well provide the position of tracer inlet, which corroborates well with the depth of a permeable zone delineated by the flowmeter. This article demonstrates the potential of nano-iron tracer test to provide the quantitative information of fracture flow paths in fractured rock.

VACUUM TRAP AND VALVE COMBINATION

DOEpatents

Milleron, N.; Levenson, L.

1963-02-19

This patent relates to a vacuum trap and valve combination suitable for use in large ultra-high vacuum systems. The vacuum trap is a chamber having an inlet and outlet opening which may be made to communicate with a chamber to be evacuated and a diffusion pump, respectively. A valve is designed to hermeticaliy seal with inlet opening and, when opened, block the line-of- sight'' between the inlet and outlet openings, while allowing a large flow path between the opened vaive and the side walls of the trap. The interior of the trap and the side of the valve facing the inlet opening are covered with an impurity absorbent, such as Zeolite or activated aluminum. Besides the advantage of combining two components of a vacuum system into one, the present invention removes the need for a baffle between the pump and the chamber to be evacuated. In one use of a specific embodiment of this invention, the transmission probability was 45 and the partial pressure of the pump fluid vapor in the vacuum chamber was at least 100 times lower than its vapor pressure. (AEC)

Observations of inner shelf cross-shore surface material transport adjacent to a coastal inlet in the northern Gulf of Mexico

NASA Astrophysics Data System (ADS)

Roth, Mathias K.; MacMahan, Jamie; Reniers, Ad; Özgökmen, Tamay M.; Woodall, Kate; Haus, Brian

2017-04-01

Motivated by the Deepwater Horizon oil spill, the Surfzone and Coastal Oil Pathways Experiment obtained Acoustic Doppler Current Profiler (ADCP) Eulerian and GPS-drifter based Lagrangian "surface" (<1 m) flow observations in the northern Gulf of Mexico to describe the influence of small-scale river plumes on surface material transport pathways in the nearshore. Lagrangian paths are qualitatively similar to surface pathlines derived from non-traditional, near-surface ADCP velocities, but both differ significantly from depth-averaged subsurface pathlines. Near-surface currents are linearly correlated with wind velocities (r =0.76 in the alongshore and r =0.85 in the cross-shore) at the 95% confidence level, and are 4-7 times larger than theoretical estimates of wind and wave-driven surface flow in an un-stratified water column. Differences in near-surface flow are attributed to the presence of a buoyant river plume forced by winds from passing extratropical storms. Plume boundary fronts induce a horizontal velocity gradient where drifters deployed outside of the plume in oceanic water routinely converge, slow, and are re-directed. When the plume flows west parallel to the beach, the seaward plume boundary front acts as a coastal barrier that prevents 100% of oceanic drifters from beaching within 27 km of the inlet. As a result, small-scale, wind-driven river plumes in the northern Gulf of Mexico act as coastal barriers that prevent offshore surface pollution from washing ashore west of river inlets.

Valve and dash-pot assembly

DOEpatents

Chang, Shih-Chih

1986-01-01

A dash-pot valve comprising a cylinder submerged in the fluid of a housing and having a piston attached to a plunger projecting into the path of closing movement of a pivotal valve member. A vortex chamber in said cylinder is provided with tangentially directed inlets to generate vortex flow upon retraction of said plunger and effect increasing resistance against said piston to progressively retard the closing rate of said valve member toward its seat.

Improved valve and dash-pot assembly

DOEpatents

Chang, S.C.

1985-04-23

A dash-pot valve comprises a cylinder submerged in the fluid of a housing and have a piston attached to a plunger projecting into the path of closing movement of a pivotal valve member. A vortex chamber in said cylinder is provided with targentially directed inlets to generate vortex flow upon retraction of said plunger and effect increasing resistance against said piston to progressively retard the closing rate of said valve member toward its seat.

Characterization of preferential flow paths between boreholes in fractured rock using a nanoscale zero-valent iron tracer test

NASA Astrophysics Data System (ADS)

Chuang, Po-Yu; Chia, Yeeping; Liou, Ya-Hsuan; Teng, Mao-Hua; Liu, Ching-Yi; Lee, Tsai-Ping

2016-11-01

Recent advances in borehole geophysical techniques have improved characterization of cross-hole fracture flow. The direct detection of preferential flow paths in fractured rock, however, remains to be resolved. In this study, a novel approach using nanoscale zero-valent iron (nZVI or `nano-iron') as a tracer was developed for detecting fracture flow paths directly. Generally, only a few rock fractures are permeable while most are much less permeable. A heat-pulse flowmeter can be used to detect changes in flow velocity for delineating permeable fracture zones in the borehole and providing the design basis for the tracer test. When nano-iron particles are released in an injection well, they can migrate through the connecting permeable fracture and be attracted to a magnet array when arriving in an observation well. Such an attraction of incoming iron nanoparticles by the magnet can provide quantitative information for locating the position of the tracer inlet. A series of field experiments were conducted in two wells in fractured rock at a hydrogeological research station in Taiwan, to test the cross-hole migration of the nano-iron tracer through permeable connected fractures. The fluid conductivity recorded in the observation well confirmed the arrival of the injected nano-iron slurry. All of the iron nanoparticles attracted to the magnet array in the observation well were found at the depth of a permeable fracture zone delineated by the flowmeter. This study has demonstrated that integrating the nano-iron tracer test with flowmeter measurement has the potential to characterize preferential flow paths in fractured rock.

Theoretical evaluation of engine auxiliary inlet design for supersonic V/STOL aircraft

NASA Technical Reports Server (NTRS)

Boles, Michael A.; Heavner, Richard L.

1988-01-01

A higher order panel method is used to evaluate the potential flow of a two dimensional supersonic V/STOL inlet. A non-symmetric analytical inlet model is developed to closely match a wind tunnel model. The analytical inlet is analyzed for flow characteristics around the lower cowl lip and auxiliary inlets. The results are obtained from the output of a computer program that is based on the Hess Panel Method which determines source strengths of panels distributed over a three dimensional body. The analytical model was designed for the implementation of drooped/translated cowl lip and auxiliary inlets as flow improvement concepts. A 40 or 70 degree droop lip can be incorporated on the inlet to determine if these geometry modifications result in flow improvements which may reduce the propensity for flow separation on the interior portion of the lip. Auxiliary inlets are employed to decrease the mass flow over the inlet lip. Thus, the peak flow velocity is reduced at the lip which also lessens the likelihood of flow separation on the interior portion of the lip. A 2, 4, and 6 inch translated lip can be employed to also decrease mass flow over the inlet lower lip in the same manner as the auxiliary inlet. The performance results of the flow improvement concepts show that three possible inlet configurations provide a situation where separation is less likely to occur. A 70 degree droop lip maintains flow conditions such that attached flow over the lower cowl lip may exist for the entire angle of attack range studied. A 0 degree droop and translated lip combination provides similar results for the angle of attack range. The third configuration consists of a 0 degree droop and auxiliary inlet combination. This configuration provides slightly less favorable results than the other two, but still allows for conditions favorable to attached flow within the inlet.

Enhancing sedimentation by improving flow conditions using parallel retrofit baffles.

PubMed

He, Cheng; Scott, Eric; Rochfort, Quintin

2015-09-01

In this study, placing parallel-connected baffles in the vicinity of the inlet was proposed to improve hydraulic conditions for enhancing TSS (total suspended solids) removal. The purpose of the retrofit baffle design is to divide the large and fast inflow into smaller and slower flows to increase flow uniformity. This avoids short-circuiting and increases residence time in the sedimentation basin. The newly proposed parallel-connected baffle configuration was assessed in the laboratory by comparing its TSS removal performance and the optimal flow residence time with those from the widely used series-connected baffles. The experimental results showed that the parallel-connected baffles outperformed the series-connected baffles because it could disperse flow faster and in less space by splitting the large inflow into many small branches instead of solely depending on flow internal friction over a longer flow path, as was the case under the series-connected baffles. Being able to dampen faster flow before entering the sedimentation basin is critical to reducing the possibility of disturbing any settled particles, especially under high inflow conditions. Also, for a large sedimentation basin, it may be more economically feasible to deploy the proposed parallel retrofit baffle in the vicinity of the inlet than series-connected baffles throughout the entire settling basin. Crown Copyright © 2015. Published by Elsevier Ltd. All rights reserved.

Inverse design of centrifugal compressor vaned diffusers in inlet shear flows

DOE Office of Scientific and Technical Information (OSTI.GOV)

Zangeneh, M.

1996-04-01

A three-dimensional inverse design method in which the blade (or vane) geometry is designed for specified distributions of circulation and blade thickness is applied to the design of centrifugal compressor vaned diffusers. Two generic diffusers are designed, one with uniform inlet flow (equivalent to a conventional design) and the other with a sheared inlet flow. The inlet shear flow effects are modeled in the design method by using the so-called ``Secondary Flow Approximation`` in which the Bernoulli surfaces are convected by the tangentially mean inviscid flow field. The difference between the vane geometry of the uniform inlet flow and nonuniformmore » inlet flow diffusers is found to be most significant from 50 percent chord to the trailing edge region. The flows through both diffusers are computed by using Denton`s three-dimensional inviscid Euler solver and Dawes` three-dimensional Navier-Stokes solver under sheared in-flow conditions. The predictions indicate improved pressure recovery and internal flow field for the diffuser designed for shear inlet flow conditions.« less

CFD simulations of the flow control performance applied for inlet of low drag high-bypass turbofan engine at cross flow regimes

NASA Astrophysics Data System (ADS)

Kursakov, I. A.; Kazhan, E. V.; Lysenkov, A. V.; Savelyev, A. A.

2016-10-01

Paper describes the optimization procedure for low cruise drag inlet of high-bypass ratio turbofan engine (HBRE). The critical cross-flow velocity when the flow separation on the lee side of the inlet channel occurs is determined. The effciency of different flow control devices used to improve the flow parameters at inlet section cross flow regime is analyzed. Boundary layer suction, bypass slot and vortex generators are considered. It is shown that flow control devices enlarge the stability range of inlet performance at cross flow regimes.

Turbine system

DOEpatents

McMahan, Kevin Weston; Dillard, Daniel Jackson

2016-05-03

A turbine system is disclosed. The turbine system includes a transition duct having an inlet, an outlet, and a passage extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of the transition duct is offset from the inlet along the longitudinal axis and the tangential axis. The turbine system further includes a turbine section connected to the transition duct. The turbine section includes a plurality of shroud blocks at least partially defining a hot gas path, a plurality of buckets at least partially disposed in the hot gas path, and a plurality of nozzles at least partially disposed in the hot gas path. At least one of a shroud block, a bucket, or a nozzle includes means for withstanding high temperatures.

Induction-heating MOCVD reactor with significantly improved heating efficiency and reduced harmful magnetic coupling

NASA Astrophysics Data System (ADS)

Li, Kuang-Hui; Alotaibi, Hamad S.; Sun, Haiding; Lin, Ronghui; Guo, Wenzhe; Torres-Castanedo, Carlos G.; Liu, Kaikai; Valdes-Galán, Sergio; Li, Xiaohang

2018-04-01

In a conventional induction-heating III-nitride metalorganic chemical vapor deposition (MOCVD) reactor, the induction coil is outside the chamber. Therefore, the magnetic field does not couple with the susceptor well, leading to compromised heating efficiency and harmful coupling with the gas inlet and thus possible overheating. Hence, the gas inlet has to be at a minimum distance away from the susceptor. Because of the elongated flow path, premature reactions can be more severe, particularly between Al- and B-containing precursors and NH3. Here, we propose a structure that can significantly improve the heating efficiency and allow the gas inlet to be closer to the susceptor. Specifically, the induction coil is designed to surround the vertical cylinder of a T-shaped susceptor comprising the cylinder and a top horizontal plate holding the wafer substrate within the reactor. Therefore, the cylinder coupled most magnetic field to serve as the thermal source for the plate. Furthermore, the plate can block and thus significantly reduce the uncoupled magnetic field above the susceptor, thereby allowing the gas inlet to be closer. The results show approximately 140% and 2.6 times increase in the heating and susceptor coupling efficiencies, respectively, as well as a 90% reduction in the harmful magnetic flux on the gas inlet.

Ultra high bypass Nacelle aerodynamics inlet flow-through high angle of attack distortion test

NASA Technical Reports Server (NTRS)

Larkin, Michael J.; Schweiger, Paul S.

1992-01-01

A flow-through inlet test program was conducted to evaluate inlet test methods and determine the impact of the fan on inlet separation when operating at large angles of attack. A total of 16 model configurations of approximately 1/6 scale were tested. A comparison of these flow-through results with powered data indicates the presence of the fan increased separation operation 3 degrees to 4 degrees over the flow through inlet. Rods and screens located at the fan face station, that redistribute the flow, achieved simulation of the powered-fan results for separation angle of attack. Concepts to reduce inlet distortion and increase angle of attack capability were also evaluated. Vortex generators located on the inlet surface increased inlet angle of attack capability up to 2 degrees and reduced inlet distortion in the separated region. Finally, a method of simulating the fan/inlet aerodynamic interaction using blockage sizing method has been defined. With this method, a static blockage device used with a flow-through model will approximate the same inlet onset of separation angle of attack and distortion pattern that would be obtained with an inlet model containing a powered fan.

Numerical Modeling of Flow Control in a Boundary-Layer-Ingesting Offset Inlet Diffuser at Transonic Mach Numbers

NASA Technical Reports Server (NTRS)

Allan, Brian G.; Owens, Lewis R.

2006-01-01

This paper will investigate the validation of the NASA developed, Reynolds-averaged Navier-Stokes (RANS) flow solver, OVERFLOW, for a boundary-layer-ingesting (BLI) offset (S-shaped) inlet in transonic flow with passive and active flow control devices as well as a baseline case. Numerical simulations are compared to wind tunnel results of a BLI inlet experiment conducted at the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel. Comparisons of inlet flow distortion, pressure recovery, and inlet wall pressures are performed. The numerical simulations are compared to the BLI inlet data at a free-stream Mach number of 0.85 and a Reynolds number of approximately 2 million based on the fanface diameter. The numerical simulations with and without tunnel walls are performed, quantifying tunnel wall effects on the BLI inlet flow. A comparison is made between the numerical simulations and the BLI inlet experiment for the baseline and VG vane cases at various inlet mass flow rates. A comparison is also made to a BLI inlet jet configuration for varying actuator mass flow rates at a fixed inlet mass flow rate. Overall, the numerical simulations were able to predict the baseline circumferential flow distortion, DPCP avg, very well within the designed operating range of the BLI inlet. A comparison of the average total pressure recovery showed that the simulations were able to predict the trends but had a negative 0.01 offset when compared to the experimental levels. Numerical simulations of the baseline inlet flow also showed good agreement with the experimental inlet centerline surface pressures. The vane case showed that the CFD predicted the correct trends in the circumferential distortion levels for varying inlet mass flow but had a distortion level that was nearly twice as large as the experiment. Comparison to circumferential distortion measurements for a 15 deg clocked 40 probe rake indicated that the circumferential distortion levels are very sensitive to the symmetry of the flow and that a misalignment of the vanes in the experiment could have resulted in this difference. The numerical simulations of the BLI inlet with jets showed good agreement with the circumferential inlet distortion levels for a range of jet actuator mass flow ratios at a fixed inlet mass flow rate. The CFD simulations for the jet case also predicted an average total pressure recovery offset that was 0.01 lower than the experiment as was seen in the baseline. Comparisons of the flow features for the jet cases revealed that the CFD predicted a much larger vortex at the engine fan-face when compare to the experiment.

Boundary-layer-ingesting inlet flow control system

NASA Technical Reports Server (NTRS)

Owens, Lewis R. (Inventor); Allan, Brian G. (Inventor)

2010-01-01

A system for reducing distortion at the aerodynamic interface plane of a boundary-layer-ingesting inlet using a combination of active and passive flow control devices is disclosed. Active flow control jets and vortex generating vanes are used in combination to reduce distortion across a range of inlet operating conditions. Together, the vortex generating vanes can reduce most of the inlet distortion and the active flow control jets can be used at a significantly reduced control jet mass flow rate to make sure the inlet distortion stays low as the inlet mass flow rate varies. Overall inlet distortion, measured and described as average SAE circumferential distortion descriptor, was maintained at a value of 0.02 or less. Advantageous arrangements and orientations of the active flow control jets and the vortex generating vanes were developed using computational fluid dynamics simulations and wind tunnel experimentations.

Water-tunnel investigation of concepts for alleviation of adverse inlet spillage interactions with external stores

NASA Technical Reports Server (NTRS)

Neuhart, Dan H.; Rhode, Matthew N.

1990-01-01

A test was conducted in the NASA Langley 16- by 24-Inch Water Tunnel to study alleviation of the adverse interactions of inlet spillage flow on the external stores of a fighter aircraft. A 1/48-scale model of a fighter aircraft was used to simulate the flow environment around the aircraft inlets and on the downstream underside of the fuselage. A controlled inlet mass flow was simulated by drawing water into the inlets. Various flow control devices were used on the underside of the aircraft model to manipulate the vortical inlet spillage flow.

Constraining Path-Dependent Processes During Basalt-CO2 Interactions with Observations From Flow-Through and Batch Experiments

NASA Astrophysics Data System (ADS)

Thomas, D.; Garing, C.; Zahasky, C.; Harrison, A. L.; Bird, D. K.; Benson, S. M.; Oelkers, E. H.; Maher, K.

2017-12-01

Predicting the timing and magnitude of CO2 storage in basaltic rocks relies partly on quantifying the dependence of reactivity on flow path and mineral distribution. Flow-through experiments that use intact cores are advantageous because the spatial heterogeneity of pore space and reactive phases is preserved. Combining aqueous geochemical analyses and petrologic characterization with non-destructive imaging techniques (e.g. micro-computed tomography) constrains the relationship between irreversible reactions, pore connectivity and accessible surface area. Our work enhances these capabilities by dynamically imaging flow through vesicular basalts with Positron Emission Tomography (PET) scanning. PET highlights the path a fluid takes by detecting photons produced during radioactive decay of an injected radiotracer (FDG). We have performed single-phase, CO2-saturated flow-through experiments with basaltic core from Iceland at CO2 sequestration conditions (50 °C; 76-90 bar Ptot). Constant flow rate and continuous pressure measurements at the inlet and outlet of the core constrain permeability. We monitor geochemical evolution through cation and anion analysis of outlet fluid sampled periodically. Before and after reaction, we perform PET scans and characterize the core using micro-CT. The PET scans indicate a discrete, localized flow path that appears to be a micro-crack connecting vesicles, suggesting that vesicle-lining minerals are immediately accessible and important reactants. Rapid increases in aqueous cation concentration, pH and HCO3- indicate that the rock reacts nearly immediately after CO2 injection. After 24 hours the solute release decreases, which may reflect a transition to reaction with phases with slower kinetic dissolution rates (e.g. zeolites and glasses to feldspar), a decrease in available reactive surface area or precipitation. We have performed batch experiments using crushed material of the same rock to elucidate the effect of flow path geometry and mineral accessibility on geochemical evolution. Interestingly, surface area-normalized dissolution rates as evinced by SiO2 release in all experiments approach similar values ( 10-15 mol/cm2/s). Our experiments show how imaging techniques are helpful in interpreting path-dependent processes in open systems.

Numerical Modeling of Flow Control in a Boundary-Layer-Ingesting Offset Inlet Diffuser at Transonic Mach Numbers

NASA Technical Reports Server (NTRS)

Allan Brian G.; Owens, Lewis, R.

2006-01-01

This paper will investigate the validation of a NASA developed, Reynolds-averaged Navier-Stokes (RANS) flow solver, OVERFLOW, for a boundary-layer-ingesting (BLI) offset (S-shaped) inlet in transonic flow with passive and active flow control devices as well as the baseline case. Numerical simulations are compared to wind tunnel results of a BLI inlet conducted at the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel. Comparisons of inlet flow distortion, pressure recovery, and inlet wall pressures are performed. The numerical simulations are compared to the BLI inlet data at a freestream Mach number of 0.85 and a Reynolds number of approximately 2 million based on the length of the fan-face diameter. The numerical simulations with and without wind tunnel walls are performed, quantifying effects of the tunnel walls on the BLI inlet flow measurements. The wind tunnel test evaluated several different combinations of jet locations and mass flow rates as well as a vortex generator (VG) vane case. The numerical simulations will be performed on a single jet configuration for varying actuator mass flow rates at a fix inlet mass flow condition. Validation of the numerical simulations for the VG vane case will also be performed for varying inlet mass flow rates. Overall, the numerical simulations were able to predict the baseline circumferential flow distortion, DPCPavg, very well for comparisons made within the designed operating range of the BLI inlet. However the CFD simulations did predict a total pressure recovery that was 0.01 lower than the experiment. Numerical simulations of the baseline inlet flow also showed good agreement with the experimental inlet centerline surface pressures. The vane case showed that the CFD predicted the correct trends in the circumferential distortion for varying inlet mass flow but had a distortion level that was nearly twice as large as the experiment. Comparison to circumferential distortion measurements for a 15 deg clocked 40 probe rake indicated that the circumferential distortion levels are very sensitive to the symmetry of the flow and that a miss alignment of the vanes in the experiment could have resulted in this difference. The numerical simulations of the BLI inlet with jets showed good agreement with the circumferential inlet distortion levels for a range of jet actuator mass flow ratios at a fixed inlet mass flow rate. The CFD simulations for the jet case also predicted an average total pressure recovery that was 0.01 lower than the experiment as was seen in the baseline. Comparison of the flow features the jet case revealed that the CFD predicted a much larger vortex at the engine fan-face when compare to the experiment.

Rotary blood pump

NASA Technical Reports Server (NTRS)

Benkowski, Robert J. (Inventor); Kiris, Cetin (Inventor); Kwak, Dochan (Inventor); Rosenbaum, Bernard J. (Inventor); Bacak, James W. (Inventor); DeBakey, Michael E. (Inventor)

1999-01-01

A blood pump that comprises a pump housing having a blood flow path therethrough, a blood inlet, and a blood outlet; a stator mounted to the pump housing, the stator having a stator field winding for producing a stator magnetic field; a flow straightener located within the pump housing, and comprising a flow straightener hub and at least one flow straightener blade attached to the flow straightener hub; a rotor mounted within the pump housing for rotation in response to the stator magnetic field, the rotor comprising an inducer and an impeller; the inducer being located downstream of the flow straightener, and comprising an inducer hub and at least one inducer blade attached to the inducer hub; the impeller being located downstream of the inducer, and comprising an impeller hub and at least one impeller blade attached to the impeller hub; and preferably also comprising a diffuser downstream of the impeller, the diffuser comprising a diffuser hub and at least one diffuser blade. Blood flow stagnation and clot formation within the pump are minimized by, among other things, providing the inducer hub with a diameter greater than the diameter of the flow straightener hub; by optimizing the axial spacing between the flow straightener hub and the inducer hub, and between the impeller hub and the diffuser hub; by optimizing the inlet angle of the diffuser blades; and by providing fillets or curved transitions between the upstream end of the inducer hub and the shaft mounted therein, and between the impeller hub and the shaft mounted therein.

Contaminate Control Device

NASA Technical Reports Server (NTRS)

Howe, Robert H. (Inventor); Flynn, Kenneth P. (Inventor); Stapleton, Thomas J. (Inventor)

2014-01-01

A contaminate control device for filtering contaminates from a gas such as air is provided. The device includes a housing having a first inlet and a first outlet. An axial flow filter is fluidly coupled between the first inlet and the first outlet, the axial flow filter has a second inlet and a second outlet. A second filter disposed about the axial flow filter and is fluidly coupled between the first inlet and the first outlet, the second filter having a third inlet on an inner diameter and a third outlet disposed on an outer diameter. A flow restrictor is fluidly coupled between the second inlet and the first inlet.

Boundary-Layer-Ingesting Inlet Flow Control

NASA Technical Reports Server (NTRS)

Owens, Lewis R.; Allan, Brian G.; Gorton, Susan A.

2006-01-01

This paper gives an overview of a research study conducted in support of the small-scale demonstration of an active flow control system for a boundary-layer-ingesting (BLI) inlet. The effectiveness of active flow control in reducing engine inlet circumferential distortion was assessed using a 2.5% scale model of a 35% boundary-layer-ingesting flush-mounted, offset, diffusing inlet. This experiment was conducted in the NASA Langley 0.3-meter Transonic Cryogenic Tunnel at flight Mach numbers with a model inlet specifically designed for this type of testing. High mass flow actuators controlled the flow through distributed control jets providing the active flow control. A vortex generator point design configuration was also tested for comparison purposes and to provide a means to examine a hybrid vortex generator and control jets configuration. Measurements were made of the onset boundary layer, the duct surface static pressures, and the mass flow through the duct and the actuators. The distortion and pressure recovery were determined by 40 total pressure measurements on 8 rake arms each separated by 45 degrees and were located at the aerodynamic interface plane. The test matrix was limited to a maximum free-stream Mach number of 0.85 with scaled mass flows through the inlet for that condition. The data show that the flow control jets alone can reduce circumferential distortion (DPCP(sub avg)) from 0.055 to about 0.015 using about 2.5% of inlet mass flow. The vortex generators also reduced the circumferential distortion from 0.055 to 0.010 near the inlet mass flow design point. Lower inlet mass flow settings with the vortex generator configuration produced higher distortion levels that were reduced to acceptable levels using a hybrid vortex generator/control jets configuration that required less than 1% of the inlet mass flow.

Boundary-Layer-Ingesting Inlet Flow Control

NASA Technical Reports Server (NTRS)

Owens, Lewis R.; Allan, Brian G.; Gorton, Susan A.

2006-01-01

This paper gives an overview of a research study conducted in support of the small-scale demonstration of an active flow control system for a boundary-layer-ingesting (BLI) inlet. The effectiveness of active flow control in reducing engine inlet circumferential distortion was assessed using a 2.5% scale model of a 35% boundary-layer-ingesting flush-mounted, offset, diffusing inlet. This experiment was conducted in the NASA Langley 0.3-meter Transonic Cryogenic Tunnel at flight Mach numbers with a model inlet specifically designed for this type of testing. High mass flow actuators controlled the flow through distributed control jets providing the active flow control. A vortex generator point design configuration was also tested for comparison purposes and to provide a means to examine a hybrid vortex generator and control jets configuration. Measurements were made of the onset boundary layer, the duct surface static pressures, and the mass flow through the duct and the actuators. The distortion and pressure recovery were determined by 40 total pressure measurements on 8 rake arms each separated by 45 degrees and were located at the aerodynamic interface plane. The test matrix was limited to a maximum free-stream Mach number of 0.85 with scaled mass flows through the inlet for that condition. The data show that the flow control jets alone can reduce circumferential distortion (DPCPavg) from 0.055 to about 0.015 using about 2.5% of inlet mass flow. The vortex generators also reduced the circumferential distortion from 0.055 to 0.010 near the inlet mass flow design point. Lower inlet mass flow settings with the vortex generator configuration produced higher distortion levels that were reduced to acceptable levels using a hybrid vortex generator/control jets configuration that required less than 1% of the inlet mass flow.

Inlet flow field investigation. Part 1: Transonic flow field survey

NASA Technical Reports Server (NTRS)

Yetter, J. A.; Salemann, V.; Sussman, M. B.

1984-01-01

A wind tunnel investigation was conducted to determine the local inlet flow field characteristics of an advanced tactical supersonic cruise airplane. A data base for the development and validation of analytical codes directed at the analysis of inlet flow fields for advanced supersonic airplanes was established. Testing was conducted at the NASA-Langley 16-foot Transonic Tunnel at freestream Mach numbers of 0.6 to 1.20 and angles of attack from 0.0 to 10.0 degrees. Inlet flow field surveys were made at locations representative of wing (upper and lower surface) and forebody mounted inlet concepts. Results are presented in the form of local inlet flow field angle of attack, sideflow angle, and Mach number contours. Wing surface pressure distributions supplement the flow field data.

Effect of inlet modelling on surface drainage in coupled urban flood simulation

NASA Astrophysics Data System (ADS)

Jang, Jiun-Huei; Chang, Tien-Hao; Chen, Wei-Bo

2018-07-01

For a highly developed urban area with complete drainage systems, flood simulation is necessary for describing the flow dynamics from rainfall, to surface runoff, and to sewer flow. In this study, a coupled flood model based on diffusion wave equations was proposed to simulate one-dimensional sewer flow and two-dimensional overland flow simultaneously. The overland flow model provides details on the rainfall-runoff process to estimate the excess runoff that enters the sewer system through street inlets for sewer flow routing. Three types of inlet modelling are considered in this study, including the manhole-based approach that ignores the street inlets by draining surface water directly into manholes, the inlet-manhole approach that drains surface water into manholes that are each connected to multiple inlets, and the inlet-node approach that drains surface water into sewer nodes that are connected to individual inlets. The simulation results were compared with a high-intensity rainstorm event that occurred in 2015 in Taipei City. In the verification of the maximum flood extent, the two approaches that considered street inlets performed considerably better than that without street inlets. When considering the aforementioned models in terms of temporal flood variation, using manholes as receivers leads to an overall inefficient draining of the surface water either by the manhole-based approach or by the inlet-manhole approach. Using the inlet-node approach is more reasonable than using the inlet-manhole approach because the inlet-node approach greatly reduces the fluctuation of the sewer water level. The inlet-node approach is more efficient in draining surface water by reducing flood volume by 13% compared with the inlet-manhole approach and by 41% compared with the manhole-based approach. The results show that inlet modeling has a strong influence on drainage efficiency in coupled flood simulation.

Effects of inlet flow field conditions on the performance of centrifugal compressor diffusers: Part 1 -- Discrete-passage diffuser

DOE Office of Scientific and Technical Information (OSTI.GOV)

Filipenco, V.G.; Deniz, S.; Johnston, J.M.

2000-01-01

This is Part 1 of a two-part paper considering the performance of radial diffusers for use in a high-performance centrifugal compressor. Part 1 reports on discrete-passage diffusers, while Part 2 describes a test of a straight-channel diffuser designed for equivalent duty. Two builds of discrete-passage diffuser were tested, with 30 and 38 separate passages. Both the 30 and 38 passage diffusers investigated showed comparable range of unstalled operation and similar level of overall diffuser pressure recovery. The paper concentrates on the influence of inlet flow conditions on the pressure recovery and operating range of radial diffusers for centrifugal compressor stages.more » The flow conditions examined include diffuser inlet Mach number, flow angle, blockage, and axial flow nonuniformity. The investigation was carried out in a specially built test facility, designed to provide a controlled inlet flow field to the test diffusers. The facility can provide a wide range of diffuser inlet velocity profile distortion and skew with Mach numbers up to unity and flow angles of 63 to 75 deg from the radical direction. The consequences of different averaging methods for the inlet total pressure distributions, which are needed in the definition of diffuser pressure recovery coefficient for nonuniform diffuser inlet conditions, were also assessed. The overall diffuser pressure recovery coefficient, based on suitably averaged inlet total pressure, was found to correlate well with the momentum-averaged flow angle into the diffuser. It is shown that the generally accepted sensitivity of diffuser pressure recovery performance to inlet flow distortion and boundary layer blockage can be largely attributed to inappropriate quantification of the average dynamic pressure at diffuser inlet. Use of an inlet dynamic pressure based on availability or mass-averaging in combination with definition of inlet flow angle based on mass average of the radial and tangential velocity at diffuser inlet removes this sensitivity.« less

High-Performance Water-Iodinating Cartridge

NASA Technical Reports Server (NTRS)

Sauer, Richard; Gibbons, Randall E.; Flanagan, David T.

1993-01-01

High-performance cartridge contains bed of crystalline iodine iodinates water to near saturation in single pass. Cartridge includes stainless-steel housing equipped with inlet and outlet for water. Bed of iodine crystals divided into layers by polytetrafluoroethylene baffles. Holes made in baffles and positioned to maximize length of flow path through layers of iodine crystals. Resulting concentration of iodine biocidal; suppresses growth of microbes in stored water or disinfects contaminated equipment. Cartridge resists corrosion and can be stored wet. Reused several times before necessary to refill with fresh iodine crystals.

The NASA Ames Hypersonic Combustor-Model Inlet CFD Simulations and Experimental Comparisons

NASA Technical Reports Server (NTRS)

Venkatapathy, E.; Tokarcik-Polsky, S.; Deiwert, G. S.; Edwards, Thomas A. (Technical Monitor)

1995-01-01

Computations have been performed on a three-dimensional inlet associated with the NASA Ames combustor model for the hypersonic propulsion experiment in the 16-inch shock tunnel. The 3-dimensional inlet was designed to have the combustor inlet flow nearly two-dimensional and of sufficient mass flow necessary for combustion. The 16-inch shock tunnel experiment is a short duration test with test time of the order of milliseconds. The flow through the inlet is in chemical non-equilibrium. Two test entries have been completed and limited experimental results for the inlet region of the combustor-model are available. A number of CFD simulations, with various levels of simplifications such as 2-D simulations, 3-D simulations with and without chemical reactions, simulations with and without turbulent conditions, etc., have been performed. These simulations have helped determine the model inlet flow characteristics and the important factors that affect the combustor inlet flow and the sensitivity of the flow field to these simplifications. In the proposed paper, CFD modeling of the hypersonic inlet, results from the simulations and comparison with available experimental results will be presented.

Incidence angle bounds for lip flow separation of three 13.97-centimeter-diameter inlets

NASA Technical Reports Server (NTRS)

Luidens, R. W.; Abbott, J. M.

1976-01-01

Low speed wind tunnel tests were conducted to establish a procedure for determining inlet-lip flow separation and to make preliminary examination of the incidence angle bounds for lip flow separation on inlets intended for the nacelles of STOL (short takeoff and landing) aircraft. Three inlets were tested. Two of the inlets had short centerbodies with lower lip area contraction ratios of 1.30 and 1.44. The third inlet had a cylindrical centerbody extended forward into the inlet throat with a lower lip area contraction ratio of 1.44. The inlets were sized to fit a 13.97 centimeter-diameter fan. For inlet throat Mach numbers less than about 0.43, the lip flow separation angle was increased by either increasing the ratio of throat velocity to freestream velocity (Vt/Vo) or by increasing the lower lip area contraction ratio. For throat Mach numbers greater than a certain value (ranging from 0.43 to 0.52), increasing throat Mach number in some cases resulted in a decrease in the lip flow separation angle. Extending a cylindrical centerbody into the inlet throat increased the flow separation angle for nearly all values of Vt/Vo.

Aerodynamic performance of 0.4066-scale model of JT8D refan stage with S-duct inlet

NASA Technical Reports Server (NTRS)

Moore, R. D.; Kovich, G.; Lewis, G. W., Jr.

1977-01-01

A scale model of the JT8D refan stage was tested with a scale model of the S-duct inlet design for the refanned Boeing 727 center engine. Detailed survey data of pressures, temperatures, and flow angles were obtained over a range of flows at speeds from 70 to 97 percent of design speed. Two S-duct configurations were tested; one with a bellmouth inlet and the other with a flight lip inlet. The results indicated that the overall performance was essentially unaffected by the distortion generated by the S-duct inlet. The stall weight flow increased by less than 0.5 kg/sec (approximately 1.5% of design flow) with the S-duct inlet compared with that obtained with uniform flow. The detailed measurements indicated that the inlet guide vane (IGV) significantly reduced circumferential variations. For example, the flow angles ahead of the IGV were positive in the right half of the inlet and negative in the left half. Behind the IGV, the flow angles tended to be more uniform circumferentially.

The effect of inlet boundary conditions in image-based CFD modeling of aortic flow

NASA Astrophysics Data System (ADS)

Madhavan, Sudharsan; Kemmerling, Erica Cherry

2016-11-01

CFD of cardiovascular flow is a growing and useful field, but simulations are subject to a number of sources of uncertainty which must be quantified. Our work focuses on the uncertainty introduced by the selection of inlet boundary conditions in an image-based, patient-specific model of the aorta. Specifically, we examined the differences between plug flow, fully developed parabolic flow, linear shear flows, skewed parabolic flow profiles, and Womersley flow. Only the shape of the inlet velocity profile was varied-all other parameters were held constant between simulations, including the physiologically realistic inlet flow rate waveform and outlet flow resistance. We found that flow solutions with different inlet conditions did not exhibit significant differences beyond 1 . 75 inlet diameters from the aortic root. Time averaged wall shear stress (TAWSS) was also calculated. The linear shear velocity boundary condition solution exhibited the highest spatially averaged TAWSS, about 2 . 5 % higher than the fully developed parabolic velocity boundary condition, which had the lowest spatially averaged TAWSS.

Effects of inlet distortion on the development of secondary flows in a subsonic axial inlet compressor rotor. Ph.D. Thesis - Toledo Univ., OH

NASA Technical Reports Server (NTRS)

Owen, Albert K.

1991-01-01

Detailed flow measurements were taken inside an isolated axial compressor rotor operating subsonically near peak efficiency. Laser anemometer measurements were made with two inlet velocity profiles. One profile consisted of an unmodified baseline flow, and the second profile was distorted by placing axisymmetric screens on the hub and shroud well upstream of the rotor. A primary flow is defined in the rotor and deviations from this primary flow for each inlet flow condition identified. A comparison between the two flow deviations is made to assess the development of a passage vortex due to the distortion of the inlet flow. A comparison of experimental results with computational predictions from a Navier-Stokes solver showed good agreement between predicted and measured flow. Measured results indicate that a distorted inlet profile has minimal effect on the development of the flow in the rotor passage and the resulting passage vortex.

Assessment of inlet efficiency through a 3D simulation: numerical and experimental comparison.

PubMed

Gómez, Manuel; Recasens, Joan; Russo, Beniamino; Martínez-Gomariz, Eduardo

2016-10-01

Inlet efficiency is a requirement for characterizing the flow transfers between surface and sewer flow during rain events. The dual drainage approach is based on the joint analysis of both upper and lower drainage levels, and the flow transfer is one of the relevant elements to define properly this joint behaviour. This paper presents the results of an experimental and numerical investigation about the inlet efficiency definition. A full scale (1:1) test platform located in the Technical University of Catalonia (UPC) reproduces both the runoff process in streets and the water entering the inlet. Data from tests performed on this platform allow the inlet efficiency to be estimated as a function of significant hydraulic and geometrical parameters. A reproduction of these tests through a numerical three-dimensional code (Flow-3D) has been carried out simulating this type of flow by solving the RANS equations. The aim of the work was to reproduce the hydraulic performance of a previously tested grated inlet under several flow and geometric conditions using Flow-3D as a virtual laboratory. This will allow inlet efficiencies to be obtained without previous experimental tests. Moreover, the 3D model allows a better understanding of the hydraulics of the flow interception and the flow patterns approaching the inlet.

An experimental description of the flow in a centrifugal compressor from alternate stall to surge

NASA Astrophysics Data System (ADS)

Moënne-Loccoz, V.; Trébinjac, I.; Benichou, E.; Goguey, S.; Paoletti, B.; Laucher, P.

2017-08-01

The present paper gives the experimental results obtained in a centrifugal compressor stage designed and built by SAFRAN Helicopter Engines. The compressor is composed of inlet guide vanes, a backswept splittered unshrouded impeller, a splittered vaned radial diffuser and axial outlet guide vanes. Previous numerical simulations revealed a particular S-shape pressure rise characteristic at partial rotation speed and predicted an alternate flow pattern in the vaned radial diffuser at low mass flow rate. This alternate flow pattern involves two adjacent vane passages. One passage exhibits very low momentum and a low pressure recovery, whereas the adjacent passage has very high momentum in the passage inlet and diffuses efficiently. Experimental measurements confirm the S-shape of the pressure rise characteristic even if the stability limit experimentally occurs at higher mass flow than numerically predicted. At low mass flow the alternate stall pattern is confirmed thanks to the data obtained by high-frequency pressure sensors. As the compressor is throttled the path to instability has been registered and a first scenario of the surge inception is given. The compressor first experiences a steady alternate stall in the diffuser. As the mass flow decreases, the alternate stall amplifies and triggers the mild surge in the vaned diffuser. An unsteady behavior results from the interaction of the alternate stall and the mild surge. Finally, when the pressure gradient becomes too strong, the alternate stall blows away and the compressor enters into deep surge.

Experimental and Computational Evaluation of Flush-Mounted, S-Duct Inlets

NASA Technical Reports Server (NTRS)

Berrier, Bobby L.; Allan, Brian G.

2004-01-01

A new high Reynolds number test capability for boundary layer ingesting inlets has been developed for the NASA Langley Research Center 0.3-Meter Transonic Cryogenic Tunnel. Using this new capability. an experimental investigation of four S-duct inlet configurations was conducted. A computational study of one of the inlets was also conducted using a Navier-Stokes solver. The objectives of this investigation were to: 1) develop a new high Reynolds number inlet test capability for flush-mounted inlets; 2) provide a database for CFD tool validation; 3) evaluate the performance of S-duct inlets with large amounts of boundary layer ingestion; and 4) provide a baseline inlet for future inlet flow-control studies. Tests were conducted at Mach numbers from 0.25 to 0.83. Reynolds numbers (based on duct exit diameter) from 5.1 million to a full-scale value of 13.9 million, and inlet mass-flow ratios from 0.39 to 1.58 depending on Mach number. Results of the experimental study indicate that inlet pressure recovery generally decreased and inlet distortion generally increased with increasing Mach number. Except at low Mach numbers, increasing inlet mass-flow increased pressure recovery and increased distortion. Increasing the amount of boundary layer ingestion or ingesting a boundary layer with a distorted profile decreased pressure recovery and increased distortion. Finally, increasing Reynolds number had almost no effect on inlet distortion but increased inlet recovery by about one-half percent at a Mach number near cruise. The computational results captured the inlet pressure recovery and distortion trends with Mach number and inlet mass-flow well: the reversal of the pressure recovery trend with increasing inlet mass-flow at low and high Mach numbers was predicted by CFD. However, CFD results were generally more pessimistic (larger losses) than measured experimentally.

Research on the performance of low-lift diving tubular pumping system by CFD and Test

NASA Astrophysics Data System (ADS)

Xia, Chenzhi; Cheng, Li; Liu, Chao; Zhou, Jiren; Tang, Fangping; Jin, Yan

2016-11-01

Post-diving tubular pump is always used in large-discharge & low-head irrigation or storm drainage pumping station, its impeller and motor share the same shaft. Considering diving tubular pump system's excellent hydraulic performance, compact structure, good noise resistance and low operating cost, it is used in Chinese pump stations. To study the hydraulic performance and pressure fluctuation of inlet and outlet passage in diving tubular pump system, both of steady and unsteady full flow fields are numerically simulated at three flow rate conditions by using CFD commercial software. The asymmetry of the longitudinal structure of inlet passage affects the flow pattern on outlet. Especially at small flow rate condition, structural asymmetry will result in the uneven velocity distribution on the outlet of passage inlet. The axial velocity distribution uniformity increases as the flow rate increases on the inlet of passage inlet, and there is a positive correlation between hydraulic loss in the passage inlet and flow rate's quadratic. The axial velocity distribution uniformity on the outlet of passage inlet is 90% at design flow rate condition. The predicted result shows the same trend with test result, and the range of high efficiency area between predicted result and test result is almost identical. The dominant frequency of pressure pulsation is low frequency in inlet passage at design condition. The dominant frequency is high frequency in inlet passage at small and large flow rate condition. At large flow rate condition, the flow pattern is significantly affected by the rotation of impeller in inlet passage. At off-design condition, the pressure pulsation is strong at outlet passage. At design condition, the dominant frequency is 35.57Hz, which is double rotation frequency.

Flow Energy Piezoelectric Bimorph Nozzle Harvester

NASA Technical Reports Server (NTRS)

Walkemeyer, Phillip E. (Inventor); Tosi, Phillipe (Inventor); Corbett, Thomas Gary (Inventor); Hall, Jeffrey L. (Inventor); Lee, Hyeong Jae (Inventor); Arrazola, Alvaro Jose (Inventor); Sherrit, Stewart (Inventor); Colonius, Tim (Inventor); Kim, Namhyo (Inventor); Sun, Kai (Inventor)

2016-01-01

A flow energy harvesting device having a harvester pipe includes a flow inlet that receives flow from a primary pipe, a flow outlet that returns the flow into the primary pipe, and a flow diverter within the harvester pipe having an inlet section coupled to the flow inlet, a flow constriction section coupled to the inlet section and positioned at a midpoint of the harvester pipe and having a spline shape with a substantially reduced flow opening size at a constriction point along the spline shape, and an outlet section coupled to the constriction section. The harvester pipe may further include a piezoelectric structure extending from the inlet section through the constriction section and point such that the fluid flow past the constriction point results in oscillatory pressure amplitude inducing vibrations in the piezoelectric structure sufficient to cause a direct piezoelectric effect and to generate electrical power for harvesting.

Effects of interstage diffuser flow distortion on the performance of a 15.41-centimeter tip diameter axial power turbine

NASA Technical Reports Server (NTRS)

Mclallin, K. L.; Kofskey, M. G.; Civinskas, K. C.

1983-01-01

The performance of a variable-area stator, axial flow power turbine was determined in a cold-air component research rig for two inlet duct configurations. The two ducts were an interstage diffuser duct and an accelerated-flow inlet duct which produced stator inlet boundary layer flow blockages of 11 percent and 3 percent, respectively. Turbine blade total efficiency at design point was measured to be 5.3 percent greater with the accelerated-flow inlet duct installed due to the reduction in inlet blockage. Blade component measurements show that of this performance improvement, 35 percent occurred in the stator and 65 percent occurred in the rotor. Analysis of inlet duct internal flow using an Axisymmetric Diffuser Duct Code (ADD Code) were in substantial agreement with the test data.

Preferential flow paths in fractured rock detected by cross-borehole nano-iron tracer test

NASA Astrophysics Data System (ADS)

Chia, Yeeping; Chuang, Po-Yu

2017-04-01

Characterization of the preferential flow paths and their hydraulic properties is desirable for developing a hydrogeological conceptual model in fractured rock. However, the heterogeneity and anisotropy of the hydraulic property often make it difficult to understand groundwater flow paths through fractures. In this study, we adopted nanoscale zero-valent iron (nZVI) as a tracer to characterize fracture connectivity and hydraulic properties. A magnet array was placed in an observation well to attract arriving nZVI particles for identifying the location of incoming tracer. This novel approach was developed for the investigation of fracture flow at a hydrogeological research station in central Taiwan. A heat-pulse flowmeter test was performed to delineate the vertical distribution of permeable fractures in two boreholes, making it possible to design a field tracer test. The nZVI slurry was released in the sealed injection well. The arrival of the slurry in the observation well was evidenced by a breakthrough curve recorded by the fluid conductivity sensor as well as the nZVI particles attracted to the magnets. The iron nanoparticles attracted to the magnets provide the quantitative criteria for locating the position of tracer inlet in the observation well. The position of the magnet attracting the maximum weight of iron nanoparticles agrees well with the depth of a permeable fracture zone delineated by the flowmeter. Besides, a conventional saline tracer test was conducted in the field, producing a similar outcome as the nZVI tracer test. Our study results indicate that the nano-iron tracer test could be a promising method for the characterization of the preferential flow paths in fractured rock.

Hypersonic Combustor Model Inlet CFD Simulations and Experimental Comparisons

NASA Technical Reports Server (NTRS)

Venkatapathy, E.; TokarcikPolsky, S.; Deiwert, G. S.; Edwards, Thomas A. (Technical Monitor)

1995-01-01

Numerous two-and three-dimensional computational simulations were performed for the inlet associated with the combustor model for the hypersonic propulsion experiment in the NASA Ames 16-Inch Shock Tunnel. The inlet was designed to produce a combustor-inlet flow that is nearly two-dimensional and of sufficient mass flow rate for large scale combustor testing. The three-dimensional simulations demonstrated that the inlet design met all the design objectives and that the inlet produced a very nearly two-dimensional combustor inflow profile. Numerous two-dimensional simulations were performed with various levels of approximations such as in the choice of chemical and physical models, as well as numerical approximations. Parametric studies were conducted to better understand and to characterize the inlet flow. Results from the two-and three-dimensional simulations were used to predict the mass flux entering the combustor and a mass flux correlation as a function of facility stagnation pressure was developed. Surface heat flux and pressure measurements were compared with the computed results and good agreement was found. The computational simulations helped determine the inlet low characteristics in the high enthalpy environment, the important parameters that affect the combustor-inlet flow, and the sensitivity of the inlet flow to various modeling assumptions.

Progress of High Efficiency Centrifugal Compressor Simulations Using TURBO

NASA Technical Reports Server (NTRS)

Kulkarni, Sameer; Beach, Timothy A.

2017-01-01

Three-dimensional, time-accurate, and phase-lagged computational fluid dynamics (CFD) simulations of the High Efficiency Centrifugal Compressor (HECC) stage were generated using the TURBO solver. Changes to the TURBO Parallel Version 4 source code were made in order to properly model the no-slip boundary condition along the spinning hub region for centrifugal impellers. A startup procedure was developed to generate a converged flow field in TURBO. This procedure initialized computations on a coarsened mesh generated by the Turbomachinery Gridding System (TGS) and relied on a method of systematically increasing wheel speed and backpressure. Baseline design-speed TURBO results generally overpredicted total pressure ratio, adiabatic efficiency, and the choking flow rate of the HECC stage as compared with the design-intent CFD results of Code Leo. Including diffuser fillet geometry in the TURBO computation resulted in a 0.6 percent reduction in the choking flow rate and led to a better match with design-intent CFD. Diffuser fillets reduced annulus cross-sectional area but also reduced corner separation, and thus blockage, in the diffuser passage. It was found that the TURBO computations are somewhat insensitive to inlet total pressure changing from the TURBO default inlet pressure of 14.7 pounds per square inch (101.35 kilopascals) down to 11.0 pounds per square inch (75.83 kilopascals), the inlet pressure of the component test. Off-design tip clearance was modeled in TURBO in two computations: one in which the blade tip geometry was trimmed by 12 mils (0.3048 millimeters), and another in which the hub flow path was moved to reflect a 12-mil axial shift in the impeller hub, creating a step at the hub. The one-dimensional results of these two computations indicate non-negligible differences between the two modeling approaches.

Microelectrical Mechanical Systems Flow Control Used to Manage Engine Face Distortion in Compact Inlet Systems

NASA Technical Reports Server (NTRS)

Anderson, Bernhard H.; Miller, Daniel N.

1999-01-01

Turbofan engine-face flow distortion is one of the most troublesome and least understood problems for designers of modern engine inlet systems. One concern is that there are numerous sources of flow-field distortion that are ingested by the inlet or generated within the inlet duct itself. Among these are: (1) flow separation at the cowl lip during in-flight maneuvering, (2) flow separation on the compression surfaces due to shock-wave/boundary layer interactions, (3) spillage of the fuselage boundary layer into the inlet duct, (4) ingestion of aircraft vortices and wakes emanating from upstream disturbances, and (5) strong secondary flow gradients and flow separation induced by wall curvature within the inlet duct itself. Most developing aircraft (including the B70, F-111, F-14, Mig-25, Tornado, and Airbus A300) have experienced one or more of these types of problems, particularly at high Mach numbers and/or extreme maneuver conditions when flow distortion at the engine face exceeded the allowable limits of the engine.

The Origin of Inlet Buzz in a Mach 1.7 Low Boom Inlet Design

NASA Technical Reports Server (NTRS)

Anderson, Bernhard H.; Weir, Lois

2014-01-01

Supersonic inlets with external compression, having a good level performance at the critical operating point, exhibit a marked instability of the flow in some subcritical operation below a critical value of the capture mass flow ratio. This takes the form of severe oscillations of the shock system, commonly known as "buzz". The underlying purpose of this study is to indicate how Detached Eddy Simulation (DES) analysis of supersonic inlets will alter how we envision unsteady inlet aerodynamics, particularly inlet buzz. Presented in this paper is a discussion regarding the physical explanation underlying inlet buzz as indicated by DES analysis. It is the normal shock wave boundary layer separation along the spike surface which reduces the capture mass flow that is the controlling mechanism which determines the onset of inlet buzz, and it is the aerodynamic characteristics of a choked nozzle that provide the feedback mechanism that sustains the buzz cycle by imposing a fixed mean corrected inlet weight flow. Comparisons between the DES analysis of the Lockheed Martin Corporation (LMCO) N+2 inlet and schlieren photographs taken during the test of the Gulfstream Large Scale Low Boom (LSLB) inlet in the NASA 8x6 ft. Supersonic Wind Tunnel (SWT) show a strong similarity both in turbulent flow field structure and shock wave formation during the buzz cycle. This demonstrates the value of DES analysis for the design and understanding of supersonic inlets.

Computational effects of inlet representation on powered hypersonic, airbreathing models

NASA Technical Reports Server (NTRS)

Huebner, Lawrence D.; Tatum, Kenneth E.

1993-01-01

Computational results are presented to illustrate the powered aftbody effects of representing the scramjet inlet on a generic hypersonic vehicle with a fairing, to divert the external flow, as compared to an operating flow-through scramjet inlet. This study is pertinent to the ground testing of hypersonic, airbreathing models employing scramjet exhaust flow simulation in typical small-scale hypersonic wind tunnels. The comparison of aftbody effects due to inlet representation is well-suited for computational study, since small model size typically precludes the ability to ingest flow into the inlet and perform exhaust simulation at the same time. Two-dimensional analysis indicates that, although flowfield differences exist for the two types of inlet representations, little, if any, difference in surface aftbody characteristics is caused by fairing over the inlet.

Tidal and subtidal exchange flows at an inlet of the Wadden Sea

NASA Astrophysics Data System (ADS)

Valle-Levinson, Arnoldo; Stanev, Emil; Badewien, Thomas H.

2018-03-01

Observations of underway velocity profiles during complete spring and neap tidal cycles were used to determine whether the spatial structures of tidal and subtidal flows at a tidal inlet in a multiple-inlet embayment are consistent with those observed at single-inlet embayments. Measurements were obtained at the Otzumer Balje, one of the multiple inlets among the East Frisian Islands of the Wadden Sea. The 1.5 km-wide inlet displayed a bathymetric profile consisting of a channel ∼15 m deep flanked by <5 m shoals. Neap tide observations spanned 36 h in the period May 11-12, 2011, while spring tide measurements exceeded 48 h from May 17 to May 19, 2011. Analysis of observations indicate that frictional effects from bathymetry molded tidal flows. Spatial distributions of semidiurnal tidal current amplitude and phase conform to those predicted by an analytical model for a basin with one inlet. Maximum semidiurnal flows appear at the surface in the channel, furthest away from bottom friction effects. Therefore, Otzumer Balje displays tidal hydrodynamics that are independent of the other inlets of the embayment. Subtidal exchange flows are laterally sheared, with residual inflow in the channel combined with outflow over shoals. The spatial distribution of these residual flows follow theoretical expectations of tidally driven flows interacting with bathymetry. Such distribution is similar to the tidal residual circulation at other inlets with only one communication to the ocean, suggesting that at subtidal scales the Otzumer Balje responds to tidal forcing independently of the other inlets.

Control of DC gas flow in a single-stage double-inlet pulse tube cooler

NASA Astrophysics Data System (ADS)

Wang, C.; Thummes, G.; Heiden, C.

The use of double-inlet mode in the pulse tube cooler opens up a possibility of DC gas flow circulating around the regenerator and pulse tube. Numerical analysis shows that effects of DC flow in a single-stage pulse tube cooler are different in some aspects from that in a 4 K pulse tube cooler. For highest cooler efficiency, DC flow should be compensated to a small value, i.e. DC flow over average AC flow at regenerator inlet should be in the range -0.0013 to +0.00016. Dual valves with reversed asymmetric geometries were used for the double-inlet bypass to control the DC flow in this paper. The experiment, performed in a single-stage double-inlet pulse tube cooler, verified that the cooler performance can be significantly improved by precisely controlling the DC flow.

Numerical investigation of flow structure and pressure pulsation in the Francis-99 turbine during startup

NASA Astrophysics Data System (ADS)

Minakov, A.; Sentyabov, A.; Platonov, D.

2017-01-01

We performed numerical simulation of flow in a laboratory model of a Francis hydroturbine at startup regimes. Numerical technique for calculating of low frequency pressure pulsations in a water turbine is based on the use of DES (k-ω Shear Stress Transport) turbulence model and the approach of “frozen rotor”. The structure of the flow behind the runner of turbine was analysed. Shows the effect of flow structure on the frequency and intensity of non-stationary processes in the flow path. Two version of the inlet boundary conditions were considered. The first one corresponded measured time dependence of the discharge. Comparison of the calculation results with the experimental data shows the considerable delay of the discharge in this calculation. Second version corresponded linear approximation of time dependence of the discharge. This calculation shows good agreement with experimental results.

Testing of the NASA Hypersonics Project Combined Cycle Engine Large Scale Inlet Mode Transition Experiment (CCE LlMX)

NASA Technical Reports Server (NTRS)

Saunders, J. D.; Stueber, T. J.; Thomas, S. R.; Suder, K. L.; Weir, L. J.; Sanders, B. W.

2012-01-01

Status on an effort to develop Turbine Based Combined Cycle (TBCC) propulsion is described. This propulsion technology can enable reliable and reusable space launch systems. TBCC propulsion offers improved performance and safety over rocket propulsion. The potential to realize aircraft-like operations and reduced maintenance are additional benefits. Among most the critical TBCC enabling technologies are: 1) mode transition from turbine to scramjet propulsion, 2) high Mach turbine engines and 3) TBCC integration. To address these TBCC challenges, the effort is centered on a propulsion mode transition experiment and includes analytical research. The test program, the Combined-Cycle Engine Large Scale Inlet Mode Transition Experiment (CCE LIMX), was conceived to integrate TBCC propulsion with proposed hypersonic vehicles. The goals address: (1) dual inlet operability and performance, (2) mode-transition sequences enabling a switch between turbine and scramjet flow paths, and (3) turbine engine transients during transition. Four test phases are planned from which a database can be used to both validate design and analysis codes and characterize operability and integration issues for TBCC propulsion. In this paper we discuss the research objectives, features of the CCE hardware and test plans, and status of the parametric inlet characterization testing which began in 2011. This effort is sponsored by the NASA Fundamental Aeronautics Hypersonics project

Performance of high mach number scramjets - Tunnel vs flight

NASA Astrophysics Data System (ADS)

Landsberg, Will O.; Wheatley, Vincent; Smart, Michael K.; Veeraragavan, Ananthanarayanan

2018-05-01

While typically analysed through ground-based impulse facilities, scramjets experience significant heating loads in flight, raising engine wall temperatures and the fuel used to cool them beyond standard laboratory conditions. Hence, the present work numerically compares an access-to-space scramjet's performance at both these conditions. The Mach 12 Rectangular-to-Elliptical Shape-Transitioning scramjet flow path is examined via three-dimensional and chemically reacting Reynolds-averaged Navier-Stokes solutions. Flight operation is modelled through 800 K and 1800 K inlet and combustor walls respectively, while fuel is injected at both inlet- and combustor-based stations at 1000 K stagnation temperature. Room temperature walls and fuel plena model shock tunnel conditions. Mixing and combustion performance indicates that while flight conditions promote rapid mixing, high combustor temperatures inhibit the completion of reaction pathways, with reactant dissociation reducing chemical heat release by 16%. However, the heated walls in flight ensured 28% less energy was absorbed by the walls. While inlet fuel injection promotes robust burning of combustor-injected fuel, premature ignition upon the inlet in flight suggests these injectors should be moved further downstream. Coupled with counteracting differences in heat release and loss to the walls, the optimal engine design for flight may differ considerably from that which gives the best performance in the tunnel.

An efficient user-oriented method for calculating compressible flow in an about three-dimensional inlets. [panel method

NASA Technical Reports Server (NTRS)

Hess, J. L.; Mack, D. P.; Stockman, N. O.

1979-01-01

A panel method is used to calculate incompressible flow about arbitrary three-dimensional inlets with or without centerbodies for four fundamental flow conditions: unit onset flows parallel to each of the coordinate axes plus static operation. The computing time is scarcely longer than for a single solution. A linear superposition of these solutions quite rigorously gives incompressible flow about the inlet for any angle of attack, angle of yaw, and mass flow rate. Compressibility is accounted for by applying a well-proven correction to the incompressible flow. Since the computing times for the combination and the compressibility correction are small, flows at a large number of inlet operating conditions are obtained rather cheaply. Geometric input is aided by an automatic generating program. A number of graphical output features are provided to aid the user, including surface streamline tracing and automatic generation of curves of curves of constant pressure, Mach number, and flow inclination at selected inlet cross sections. The inlet method and use of the program are described. Illustrative results are presented.

Three dimensional viscous analysis of a hypersonic inlet

NASA Technical Reports Server (NTRS)

Reddy, D. R.; Smith, G. E.; Liou, M.-F.; Benson, Thomas J.

1989-01-01

The flow fields in supersonic/hypersonic inlets are currently being studied at NASA Lewis Research Center using 2- and 3-D full Navier-Stokes and Parabolized Navier-Stokes solvers. These tools have been used to analyze the flow through the McDonnell Douglas Option 2 inlet which has been tested at Calspan in support of the National Aerospace Plane Program. Comparisons between the computational and experimental results are presented. These comparisons lead to better overall understanding of the complex flows present in this class of inlets. The aspects of the flow field emphasized in this work are the 3-D effects, the transition from laminar to turbulent flow, and the strong nonuniformities generated within the inlet.

Boundary-Layer-Ingesting Inlet Flow Control

NASA Technical Reports Server (NTRS)

Owens, Lewis R.; Allan, Brian G.; Gorton, Susan A.

2008-01-01

An experimental study was conducted to provide the first demonstration of an active flow control system for a flush-mounted inlet with significant boundary-layer-ingestion in transonic flow conditions. The effectiveness of the flow control in reducing the circumferential distortion at the engine fan-face location was assessed using a 2.5%-scale model of a boundary-layer-ingesting offset diffusing inlet. The inlet was flush mounted to the tunnel wall and ingested a large boundary layer with a boundary-layer-to-inlet height ratio of 35%. Different jet distribution patterns and jet mass flow rates were used in the inlet to control distortion. A vane configuration was also tested. Finally a hybrid vane/jet configuration was tested leveraging strengths of both types of devices. Measurements were made of the onset boundary layer, the duct surface static pressures, and the mass flow rates through the duct and the flow control actuators. The distortion and pressure recovery were measured at the aerodynamic interface plane. The data show that control jets and vanes reduce circumferential distortion to acceptable levels. The point-design vane configuration produced higher distortion levels at off-design settings. The hybrid vane/jet flow control configuration reduced the off-design distortion levels to acceptable ones and used less than 0.5% of the inlet mass flow to supply the jets.

Aerodynamic Design of a Dual-Flow Mach 7 Hypersonic Inlet System for a Turbine-Based Combined-Cycle Hypersonic Propulsion System

NASA Technical Reports Server (NTRS)

Sanders, Bobby W.; Weir, Lois J.

2008-01-01

A new hypersonic inlet for a turbine-based combined-cycle (TBCC) engine has been designed. This split-flow inlet is designed to provide flow to an over-under propulsion system with turbofan and dual-mode scramjet engines for flight from takeoff to Mach 7. It utilizes a variable-geometry ramp, high-speed cowl lip rotation, and a rotating low-speed cowl that serves as a splitter to divide the flow between the low-speed turbofan and the high-speed scramjet and to isolate the turbofan at high Mach numbers. The low-speed inlet was designed for Mach 4, the maximum mode transition Mach number. Integration of the Mach 4 inlet into the Mach 7 inlet imposed significant constraints on the low-speed inlet design, including a large amount of internal compression. The inlet design was used to develop mechanical designs for two inlet mode transition test models: small-scale (IMX) and large-scale (LIMX) research models. The large-scale model is designed to facilitate multi-phase testing including inlet mode transition and inlet performance assessment, controls development, and integrated systems testing with turbofan and scramjet engines.

HF radar transmissions that deviate from great-circle paths: new insight from e-POP RRI

NASA Astrophysics Data System (ADS)

Perry, G. W.; Miller, E. S.; James, H. G.; Howarth, A. D.; St-Maurice, J. P.; Yau, A. W.

2016-12-01

Significant deviations of SuperDARN radar transmissions from their expected great-circle paths have been detected at ionospheric altitudes using the Radio Receiver Instrument (RRI) on the Enhanced Polar Outflow Probe (e-POP). Experiments between SuperDARN Rankin Inlet and e-POP RRI were conducted at similar local times over consecutive days. Customized experiment modes which incorporated the agile frequency switching capabilities of each system were used. The RRI measurements show deviations of radar transmissions from their expected paths by as much as 2 or 3 SuperDARN beam widths, equivalent to 6° - 10° in bearing from Rankin Inlet. The deviations displayed a dependence on the radar carrier frequency and a day-to-day variability, suggesting that the deviations were transient in nature. We will discuss the deviations in the context of 3D ray trace modeling and measurements from the Resolute Bay Incoherent Scatter Radar - North (RISR-N). The latter provided diagnostic information of the ionosphere along the ray path between RRI and Rankin Inlet during the experiments.

A comparison of predicted and measured inlet distortion flows in a subsonic axial inlet flow compressor rotor

NASA Technical Reports Server (NTRS)

Owen, Albert K.

1992-01-01

Detailed flow measurements were taken inside an isolated axial compressor rotor operating subsonically near peak efficiency. These Laser Anemometer measurements were made with two inlet velocity profiles. One profile consisted of an unmodified baseline flow, and the second profile was distorted by placing axisymmetric screens on the hub and shroud well upstream of the rotor. A detailed comparison in the rotor relative reference frame between a Navier-Stokes solver and the measured experimental results showed good agreement between the predicted and measured flows. A primary flow is defined in the rotor and deviations and the computed predictions is made to assess the development of a passage vortex due to the distortion of the inlet flow. Computer predictions indicate that a distorted inlet profile has a minimal effect on the development of the flow in the rotor passage and the resulting passage vortex.

Eddy Current Minimizing Flow Plug for Use in Flow Conditioning and Flow Metering

NASA Technical Reports Server (NTRS)

England, John Dwight (Inventor); Kelley, Anthony R. (Inventor)

2015-01-01

An eddy-current-minimizing flow plug has open flow channels formed between the plug's inlet and outlet. Each open flow channel includes (i) a first portion that originates at the inlet face and converges to a location within the plug that is downstream of the inlet, and (ii) a second portion that originates within the plug and diverges to the outlet. The diverging second portion is approximately twice the length of the converging first portion. The plug is devoid of planar surface regions at its inlet and outlet, and in fluid flow planes of the plug that are perpendicular to the given direction of a fluid flowing therethrough.

Effects of selected design variables on three ramp, external compression inlet performance. [boundary layer control bypasses, and mass flow rate

NASA Technical Reports Server (NTRS)

Kamman, J. H.; Hall, C. L.

1975-01-01

Two inlet performance tests and one inlet/airframe drag test were conducted in 1969 at the NASA-Ames Research Center. The basic inlet system was two-dimensional, three ramp (overhead), external compression, with variable capture area. The data from these tests were analyzed to show the effects of selected design variables on the performance of this type of inlet system. The inlet design variables investigated include inlet bleed, bypass, operating mass flow ratio, inlet geometry, and variable capture area.

Inlet Flow Control and Prediction Technologies for Embedded Propulsion Systems

NASA Technical Reports Server (NTRS)

McMillan, Michelle L.; Mackie, Scott A.; Gissen, Abe; Vukasinovic, Bojan; Lakebrink, Matthew T.; Glezer, Ari; Mani, Mori; Mace, James L.

2011-01-01

Fail-safe, hybrid, flow control (HFC) is a promising technology for meeting high-speed cruise efficiency, low-noise signature, and reduced fuel-burn goals for future, Hybrid-Wing-Body (HWB) aircraft with embedded engines. This report details the development of HFC technology that enables improved inlet performance in HWB vehicles with highly integrated inlets and embedded engines without adversely affecting vehicle performance. In addition, new test techniques for evaluating Boundary-Layer-Ingesting (BLI)-inlet flow-control technologies developed and demonstrated through this program are documented, including the ability to generate a BLI-like inlet-entrance flow in a direct-connect, wind-tunnel facility, as well as, the use of D-optimal, statistically designed experiments to optimize test efficiency and enable interpretation of results. Validated improvements in numerical analysis tools and methods accomplished through this program are also documented, including Reynolds-Averaged Navier-Stokes CFD simulations of steady-state flow physics for baseline, BLI-inlet diffuser flow, as well as, that created by flow-control devices. Finally, numerical methods were employed in a ground-breaking attempt to directly simulate dynamic distortion. The advances in inlet technologies and prediction tools will help to meet and exceed "N+2" project goals for future HWB aircraft.

Vaporization chambers and associated methods

DOEpatents

Turner, Terry D.; Wilding, Bruce M.; McKellar, Michael G.; Shunn, Lee P.

2017-02-21

A vaporization chamber may include at least one conduit and a shell. The at least one conduit may have an inlet at a first end, an outlet at a second end and a flow path therebetween. The shell may surround a portion of each conduit and define a chamber surrounding the portion of each conduit. Additionally, a plurality of discrete apertures may be positioned at longitudinal intervals in a wall of each conduit, each discrete aperture of the plurality of discrete apertures sized and configured to direct a jet of fluid into each conduit from the chamber. A liquid may be vaporized by directing a first fluid comprising a liquid into the inlet at the first end of each conduit, directing jets of a second fluid into each conduit from the chamber through discrete apertures in a wall of each conduit and transferring heat from the second fluid to the first fluid.

PRELIMINARY DESIGN ANALYSIS OF AXIAL FLOW TURBINES

NASA Technical Reports Server (NTRS)

Glassman, A. J.

1994-01-01

A computer program has been developed for the preliminary design analysis of axial-flow turbines. Rapid approximate generalized procedures requiring minimum input are used to provide turbine overall geometry and performance adequate for screening studies. The computations are based on mean-diameter flow properties and a stage-average velocity diagram. Gas properties are assumed constant throughout the turbine. For any given turbine, all stages, except the first, are specified to have the same shape velocity diagram. The first stage differs only in the value of inlet flow angle. The velocity diagram shape depends upon the stage work factor value and the specified type of velocity diagram. Velocity diagrams can be specified as symmetrical, zero exit swirl, or impulse; or by inputting stage swirl split. Exit turning vanes can be included in the design. The 1991 update includes a generalized velocity diagram, a more flexible meanline path, a reheat model, a radial component of velocity, and a computation of free-vortex hub and tip velocity diagrams. Also, a loss-coefficient calibration was performed to provide recommended values for airbreathing engine turbines. Input design requirements include power or pressure ratio, mass flow rate, inlet temperature and pressure, and rotative speed. The design variables include inlet and exit diameters, stator angle or exit radius ratio, and number of stages. Gas properties are input as gas constant, specific heat ratio, and viscosity. The program output includes inlet and exit annulus dimensions, exit temperature and pressure, total and static efficiencies, flow angles, blading angles, and last stage absolute and relative Mach numbers. This program is written in FORTRAN 77 and can be ported to any computer with a standard FORTRAN compiler which supports NAMELIST. It was originally developed on an IBM 7000 series computer running VM and has been implemented on IBM PC computers and compatibles running MS-DOS under Lahey FORTRAN, and DEC VAX series computers running VMS. Format statements in the code may need to be rewritten depending on your FORTRAN compiler. The source code and sample data are available on a 5.25 inch 360K MS-DOS format diskette. This program was developed in 1972 and was last updated in 1991. IBM and IBM PC are registered trademarks of International Business Machines. MS-DOS is a registered trademark of Microsoft Corporation. DEC VAX, and VMS are trademarks of Digital Equipment Corporation.

Cooling Performance Analysis of ThePrimary Cooling System ReactorTRIGA-2000Bandung

NASA Astrophysics Data System (ADS)

Irianto, I. D.; Dibyo, S.; Bakhri, S.; Sunaryo, G. R.

2018-02-01

The conversion of reactor fuel type will affect the heat transfer process resulting from the reactor core to the cooling system. This conversion resulted in changes to the cooling system performance and parameters of operation and design of key components of the reactor coolant system, especially the primary cooling system. The calculation of the operating parameters of the primary cooling system of the reactor TRIGA 2000 Bandung is done using ChemCad Package 6.1.4. The calculation of the operating parameters of the cooling system is based on mass and energy balance in each coolant flow path and unit components. Output calculation is the temperature, pressure and flow rate of the coolant used in the cooling process. The results of a simulation of the performance of the primary cooling system indicate that if the primary cooling system operates with a single pump or coolant mass flow rate of 60 kg/s, it will obtain the reactor inlet and outlet temperature respectively 32.2 °C and 40.2 °C. But if it operates with two pumps with a capacity of 75% or coolant mass flow rate of 90 kg/s, the obtained reactor inlet, and outlet temperature respectively 32.9 °C and 38.2 °C. Both models are qualified as a primary coolant for the primary coolant temperature is still below the permitted limit is 49.0 °C.

An Interactive, Design and Educational Tool for Supersonic External-Compression Inlets

NASA Technical Reports Server (NTRS)

Benson, Thomas J.

1994-01-01

A workstation-based interactive design tool called VU-INLET was developed for the inviscid flow in rectangular, supersonic, external-compression inlets. VU-INLET solves for the flow conditions from free stream, through the supersonic compression ramps, across the terminal normal shock region and the subsonic diffuser to the engine face. It calculates the shock locations, the capture streamtube, and the additive drag of the inlet. The inlet geometry can be modified using a graphical user interface and the new flow conditions recalculated interactively. Free stream conditions and engine airflow can also be interactively varied and off-design performance evaluated. Flow results from VU-INLET can be saved to a file for a permanent record, and a series of help screens make the simulator easy to learn and use. This paper will detail the underlying assumptions of the models and the numerical methods used in the simulator.

Computation of inlet reference plane flow-field for a subscale free-jet forebody/inlet model and comparison to experimental data

NASA Astrophysics Data System (ADS)

McClure, M. D.; Sirbaugh, J. R.

1991-02-01

The computational fluid dynamics (CFD) computer code PARC3D was used to predict the inlet reference plane (IRP) flow field for a side-mounted inlet and forebody simulator in a free jet for five different flow conditions. The calculations were performed for free-jet conditions, mass flow rates, and inlet configurations that matched the free-jet test conditions. In addition, viscous terms were included in the main flow so that the viscous free-jet shear layers emanating from the free-jet nozzle exit were modeled. A measure of the predicted accuracy was determined as a function of free-stream Mach number, angle-of-attack, and sideslip angle.

Optimal Flow Control Design

NASA Technical Reports Server (NTRS)

Allan, Brian; Owens, Lewis

2010-01-01

In support of the Blended-Wing-Body aircraft concept, a new flow control hybrid vane/jet design has been developed for use in a boundary-layer-ingesting (BLI) offset inlet in transonic flows. This inlet flow control is designed to minimize the engine fan-face distortion levels and the first five Fourier harmonic half amplitudes while maximizing the inlet pressure recovery. This concept represents a potentially enabling technology for quieter and more environmentally friendly transport aircraft. An optimum vane design was found by minimizing the engine fan-face distortion, DC60, and the first five Fourier harmonic half amplitudes, while maximizing the total pressure recovery. The optimal vane design was then used in a BLI inlet wind tunnel experiment at NASA Langley's 0.3-meter transonic cryogenic tunnel. The experimental results demonstrated an 80-percent decrease in DPCPavg, the reduction in the circumferential distortion levels, at an inlet mass flow rate corresponding to the middle of the operational range at the cruise condition. Even though the vanes were designed at a single inlet mass flow rate, they performed very well over the entire inlet mass flow range tested in the wind tunnel experiment with the addition of a small amount of jet flow control. While the circumferential distortion was decreased, the radial distortion on the outer rings at the aerodynamic interface plane (AIP) increased. This was a result of the large boundary layer being distributed from the bottom of the AIP in the baseline case to the outer edges of the AIP when using the vortex generator (VG) vane flow control. Experimental results, as already mentioned, showed an 80-percent reduction of DPCPavg, the circumferential distortion level at the engine fan-face. The hybrid approach leverages strengths of vane and jet flow control devices, increasing inlet performance over a broader operational range with significant reduction in mass flow requirements. Minimal distortion level requirements are met using vanes alone, avoiding engine stall and increasing robustness of this hybrid inlet flow control approach. This design applies to aerospace applications needing flush-mounted boundary-layer-ingesting inlets.

Potential and viscous flow in VTOL, STOL or CTOL propulsion system inlets

NASA Technical Reports Server (NTRS)

Stockman, N. O.

1975-01-01

A method was developed for analyzing the flow in subsonic axisymmetric inlets at arbitrary conditions of freestream velocity, incidence angle, and inlet mass flow. An improved version of the method is discussed and comparisons of results obtained with the original and improved methods are given. Comparisons with experiments are also presented for several inlet configurations and for various conditions of the boundary layer from insignificant to separated. Applications of the method are discussed, with several examples given for specific cases involving inlets for VTOL lift fans and for STOL engine nacelles.

Experimental study of flow distribution and pressure loss with circumferential inlet and outlet manifolds

NASA Technical Reports Server (NTRS)

Dittrich, R. T.

1972-01-01

Water flow tests with circumferential inlet and outlet manifolds were conducted to determine factors affecting fluid distribution and pressure losses. Various orifice sizes and manifold geometries were tested over a range of flow velocities. With inlet manifolds, flow distribution was related directly to orifice discharge coefficients. A correlation indicated that nonuniform distribution resulted when the velocity head ratio at the orifice was not in the range of constant discharge coefficient. With outlet manifolds, nonuniform flow was related to static pressure variations along the manifold. Outlet manifolds had appreciably greater pressure losses than comparable inlet manifolds.

Characteristics Air Flow in Room Chamber Test Refrigerator Household Energy Consumption with Inlet Flow Variation

NASA Astrophysics Data System (ADS)

Susanto, Edy; Idrus Alhamid, M.; Nasruddin; Budihardjo

2018-03-01

Room Chamber is the most important in making a good Testing Laboratory. In this study, the 2-D modeling conducted to assess the effect placed the inlet on designing a test chamber room energy consumption of household refrigerators. Where the geometry room chamber is rectangular and approaching the enclosure conditions. Inlet varied over the side parallel to the outlet and compared to the inlet where the bottom is made. The purpose of this study was to determine and define the characteristics of the airflow in the room chamber using CFD simulation. CFD method is used to obtain flow characteristics in detail, in the form of vector flow velocity and temperature distribution inside the chamber room. The result found that the position of the inlet parallel to the outlet causes air flow cannot move freely to the side of the floor, even flow of air moves up toward the outlet. While by making the inlet is below, the air can move freely from the bottom up to the side of the chamber room wall as well as to help uniform flow.

Experimental Research on Optimizing Inlet Airflow of Wet Cooling Towers under Crosswind Conditions

NASA Astrophysics Data System (ADS)

Chen, You Liang; Shi, Yong Feng; Hao, Jian Gang; Chang, Hao; Sun, Feng Zhong

2018-01-01

A new approach of installing air deflectors around tower inlet circumferentially was proposed to optimize the inlet airflow and reduce the adverse effect of crosswinds on the thermal performance of natural draft wet cooling towers (NDWCT). And inlet airflow uniformity coefficient was defined to analyze the uniformity of circumferential inlet airflow quantitatively. Then the effect of air deflectors on the NDWCT performance was investigated experimentally. By contrast between inlet air flow rate and cooling efficiency, it has been found that crosswinds not only decrease the inlet air flow rate, but also reduce the uniformity of inlet airflow, which reduce NDWCT performance jointly. After installing air deflectors, the inlet air flow rate and uniformity coefficient increase, the uniformity of heat and mass transfer increases correspondingly, which improve the cooling performance. In addition, analysis on Lewis factor demonstrates that the inlet airflow optimization has more enhancement of heat transfer than mass transfer, but leads to more water evaporation loss.

Cloud-Droplet Ingestion in Engine Inlets with Inlet Velocity Ratios of 1.0 and 0.7

NASA Technical Reports Server (NTRS)

Brun, Rinaldo J

1957-01-01

The paths of cloud droplets into two engine inlets have been calculated for a wide range of meteorological and flight conditions. The amount of water in droplet form ingested by the inlets and the amount and distribution of water impinging on the inlet walls are obtained from these droplet-trajectory calculations. In both types of inlet, a prolate ellipsoid of revolution represents either part or all of the forebody at the center of an annular inlet to an engine. The configurations can also represent a fuselage of an airplane with side ram-scoop inlets. The studies were made at an angle of attack of 0 degree. The principal difference between the two inlets studied is that the inlet-air velocity of one is 0.7 that of the other. The studies of the two velocity ratios lead to some important general concepts of water ingestion in inlets.

High speed variable delivery helical screw compressor/expander automotive air conditioning and waste heat energy recovery system

DOE Office of Scientific and Technical Information (OSTI.GOV)

Gagnon, J.A.; Schaefer, D.D.; Shaw, D.N.

1980-09-02

A compact, helical screw compressor/expander unit is described that is mounted in a vehicle and connected to the vehicle engine driven drive shaft has inlet and outlet ports and a capacity control slide valve and a pressure matching or volume ratio slide valve, respectively, for said ports. A refrigerant loop includes the compressor, a condenser mounted in the path of air flow over the engine and an evaporator mounted in a fresh air/cab return air flow duct for the occupant. Heat pipes thermally connect the cab air flow duct to the engine exhaust system which also bears the vapor boiler.more » Selectively operated damper valves control the fresh air/cab return air for passage selectively over the evaporator coil and the heat pipes as well as the exhaust gas flow over opposite ends of the heat pipes and the vapor boiler.« less

A Numerical Study of Hypersonic Forebody/Inlet Integration Problem

NASA Technical Reports Server (NTRS)

Kumar, Ajay

1991-01-01

A numerical study of hypersonic forebody/inlet integration problem is presented in the form of the view-graphs. The following topics are covered: physical/chemical modeling; solution procedure; flow conditions; mass flow rate at inlet face; heating and skin friction loads; 3-D forebogy/inlet integration model; and sensitivity studies.

An Improved Model for a Once-through Counter-Cross-Flow Waste Heat Recovery Unit

DTIC Science & Technology

1983-09-01

RAnkine Cycle Energy Recovery (RACER) system. As conceived, the RACER system will be an unfired waste heat recovery system designed to convert waste... heater to arrive at the feedwater inlet. For the given geometry and flow conditions, the model will calcu- late the water inlet temperature consistent...when given feedwater inlet temperature, steam outlet tempera- ture, operating pressure, inlet and outlet gas conditions and gas flow rate. In this

Theoretical and experimental study of flow-control devices for inlets of indraft wind tunnels

NASA Technical Reports Server (NTRS)

Ross, James C.

1989-01-01

The design of closed circuit wind tunnels has historically been performed using rule of thumb which have evolved over the years into a body of useful guidelines. The development of indraft wind tunnels, however, has not been as well documented. The design of indraft wind tunnels is therefore generally performed using a more intuitive approach, often resulting in a facility with disappointing flow quality. The primary problem is a lack of understanding of the flow in the inlet as it passes through the required antiturbulence treatment. For wind tunnels which employ large contraction ratio inlets, this lack of understanding is not serious since the relatively low velocity of the flow through the inlet treatment reduces the sensitivity to improper inlet design. When designing a small contraction ratio inlet, much more careful design is needed in order to reduce the flow distortions generated by the inlet treatment. As part of the National Full Scale Aerodynamics Complex Modification Project, 2-D computational methods were developed which account for the effect of both inlet screens and guide vanes on the test section velocity distribution. Comparisons with experimental data are presented which indicate that the methods accurately compute the flow distortions generated by a screen in a nonuniform velocity field. The use of inlet guide vanes to eliminate the screen induced distortion is also demonstrated both computationally and experimentally. Extensions of the results to 3-D is demonstrated and a successful wind tunnel design is presented.

Single-stage experimental evaluation of tandem-airfoil rotor stator blading for compressors. Part 6: Data and performance for stage D

NASA Technical Reports Server (NTRS)

Clemmons, D. R.

1973-01-01

An axial flow compressor stage, having single-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 had 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 tandem-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. With uniform inlet flow, the rotor achieved a maximum adiabatic efficiency of 90.1% at design equivalent rotor speed and a pressure ratio of 1.281. The stage maximum adiabatic efficiency at design equivalent rotor speed with uniform inlet flow was 86.1% at a pressure ratio of 1.266. Hub radial, tip radial, and circumferential distortion of the inlet flow caused reductions in surge pressure ratio of approximately 2, 10 and 5%, respectively, at design rotor speed.

Simulation of two-dimensional adjustable liquid gradient refractive index (L-GRIN) microlens

NASA Astrophysics Data System (ADS)

Le, Zichun; Wu, Xiang; Sun, Yunli; Du, Ying

2017-07-01

In this paper, a two-dimensional liquid gradient refractive index (L-GRIN) microlens is designed which can be used in adjusting focusing direction and focal spot of light beam. Finite element method (FEM) is used to simulate the convection diffusion process happening in core inlet flow and cladding inlet flow. And the ray tracing method shows us the light beam focusing effect including the extrapolation of focal length and output beam spot size. When the flow rates of the core and cladding fluids are held the same between the internal and external, left and right, and upper and lower inlets, the focal length varied from 313 μm to 53.3 μm while the flow rate of liquids ranges from 500 pL/s to 10,000 pL/s. While the core flow rate is bigger than the cladding inlet flow rate, the light beam will focus on a light spot with a tunable size. By adjusting the ratio of cladding inlet flow rate including Qright/Qleft and Qup/Qdown, we get the adjustable two-dimensional focus direction rather than the one-dimensional focusing. In summary, by adjusting the flow rate of core inlet and cladding inlet, the focal length, output beam spot and focusing direction of the input light beam can be manipulated. We suppose this kind of flexible microlens can be used in integrated optics and lab-on-a-chip system.

Numerical Simulation of Boundary Layer Ingesting (BLI) Inlet-Fan Interaction

NASA Technical Reports Server (NTRS)

Giuliani, James; Chen, Jen-Ping; Beach, Timothy; Bakhle, Milind

2014-01-01

Future civil transport designs may incorporate engine inlets integrated into the body of the aircraft to take advantage of efficiency increases due to weight and drag reduction. Additional increases in engine efficiency are predicted if the inlet ingests the lower momentum boundary layer flow. Previous studies have shown, however, that efficiency benefits of Boundary Layer Ingesting (BLI) ingestion are very sensitive to the magnitude of fan and duct losses, and blade structural response to the non-uniform flow field that results from a BLI inlet has not been studied in-depth. This paper presents an effort to extend the modeling capabilities of an existing rotating turbomachinery unsteady analysis code to include the ability to solve the external and internal flow fields of a BLI inlet. The TURBO code has been a successful tool in evaluating fan response to flow distortions for traditional engine/inlet integrations, such as the development of rotating stall and inlet distortion through compressor stages. This paper describes the first phase of an effort to extend the TURBO model to calculate the external and inlet flowfield upstream of fan so that accurate pressure distortions that result from BLI configurations can be computed and used to analyze fan aerodynamics and structural response. To validate the TURBO program modifications for the BLI flowfield, experimental test data obtained by NASA for a flushmounted S-duct with large amounts of boundary layer ingestion was modeled. Results for the flow upstream and in the inlet are presented and compared to experimental data for several high Reynolds number flows to validate the modifications to the solver. Quantitative data is presented that indicates good predictive capability of the model in the upstream flow. A representative fan is attached to the inlet and results are presented for the coupled inlet/fan model. The impact on the total pressure distortion at the AIP after the fan is attached is examined.

System and method for determining coolant level and flow velocity in a nuclear reactor

DOEpatents

Brisson, Bruce William; Morris, William Guy; Zheng, Danian; Monk, David James; Fang, Biao; Surman, Cheryl Margaret; Anderson, David Deloyd

2013-09-10

A boiling water reactor includes a reactor pressure vessel having a feedwater inlet for the introduction of recycled steam condensate and/or makeup coolant into the vessel, and a steam outlet for the discharge of produced steam for appropriate work. A fuel core is located within a lower area of the pressure vessel. The fuel core is surrounded by a core shroud spaced inward from the wall of the pressure vessel to provide an annular downcomer forming a coolant flow path between the vessel wall and the core shroud. A probe system that includes a combination of conductivity/resistivity probes and/or one or more time-domain reflectometer (TDR) probes is at least partially located within the downcomer. The probe system measures the coolant level and flow velocity within the downcomer.

Gas turbine combustor exit piece with hinged connections

DOE Office of Scientific and Technical Information (OSTI.GOV)

Charron, Richard C.; Pankey, William W.

2016-04-26

An exit piece (66) with an inlet throat (67) that conducts a combustion gas flow (36A) in a path (82) from a combustor (63) to an annular chamber (68) that feeds the first blade section (37) of a gas turbine (26). The exit piece further includes an outlet portion (69) that forms a circumferential segment of the annular chamber. The outlet portion interconnects with adjacent outlet portions by hinges (78A, 78B, 80A, 80B). Each hinge may have a hinge axis (82A, 82B) parallel to a centerline (21) of the turbine. Respective gas flows (36A) are configured by an assembly (60)more » of the exit pieces to converge on the feed chamber (68) into a uniform helical flow that drives the first blade section with minimal circumferential variations in force.« less

Evaluation and application of the Baldwin-Lomax turbulence model in two-dimensional, unsteady, compressible boundary layers with and without separation in engine inlets

NASA Technical Reports Server (NTRS)

Sakowski, Barbara; Darling, Douglas; Roach, Robert L.; Vandewall, Allan

1992-01-01

There is a practical need to model high speed flows that exist in jet engine inlets. The boundary layers that form in these inlets may be turbulent or laminar and either separated or attached. Also, unsteady supersonic inlets may be subject to frequent changes in operating conditions. Some changes in the operating conditions of the inlets may include varying the inlet geometry, bleeds and bypasses, and rotating or translating the centerbody. In addition, the inlet may be either started or unstarted. Therefore, a CFD code, used to model these inlets, may have to run for several different cases. Also, since the flow conditions through an unsteady inlet may be continually fluctuating, the CFD code which models these flows may have to be run over many time steps. Therefore, it would be beneficial that the code run quickly. Many turbulence models, however, are cumbersome to implement and require a lot of computer time to run, since they add to the number of differential equations to be solved to model a flow. The Baldwin-Lomax turbulence model is a popular model. It is an algebraic, eddy viscosity model. The Baldwin-Lomax model is used in many CFD codes because it is quick and easy to implement. In this paper, we will discuss implementing the Baldwin-Lomax turbulence model for both steady and unsteady compressible flows. In addition, these flows may be either separated or attached. In order to apply this turbulence model to flows which may be subjected to these conditions, certain modifications should be made to the original Baldwin-Lomax model. We will discuss these modifications and determine whether the Baldwin-Lomax model is a viable turbulence model that produces reasonably accurate results for high speed flows that can be found in engine inlets.

Unstart coupling mechanism analysis of multiple-modules hypersonic inlet.

PubMed

Hu, Jichao; Chang, Juntao; Wang, Lei; Cao, Shibin; Bao, Wen

2013-01-01

The combination of multiplemodules in parallel manner is an important way to achieve the much higher thrust of scramjet engine. For the multiple-modules scramjet engine, when inlet unstarted oscillatory flow appears in a single-module engine due to high backpressure, how to interact with each module by massflow spillage, and whether inlet unstart occurs in other modules are important issues. The unstarted flowfield and coupling characteristic for a three-module hypersonic inlet caused by center module II and side module III were, conducted respectively. The results indicate that the other two hypersonic inlets are forced into unstarted flow when unstarted phenomenon appears on a single-module hypersonic inlet due to high backpressure, and the reversed flow in the isolator dominates the formation, expansion, shrinkage, and disappearance of the vortexes, and thus, it is the major factor of unstart coupling of multiple-modules hypersonic inlet. The coupling effect among multiple modules makes hypersonic inlet be more likely unstarted.

Preliminary Results of the Determination of Inlet-Pressure Distortion Effects on Compressor Stall and Altitude Operating Limits of the J57-P-1 Turbojet Engine

NASA Technical Reports Server (NTRS)

Wallner, L. E.; Lubick, R. J.; Chelko, L. J.

1955-01-01

During an investigation of the J57-P-1 turbojet engine in the Lewis altitude wind tunnel, effects of inlet-flow distortion on engine stall characteristics and operating limits were determined. In addition to a uniform inlet-flow profile, the inlet-pressure distortions imposed included two radial, two circumferential, and one combined radial-circumferential profile. Data were obtained over a range of compressor speeds at an altitude of 50,000 and a flight Mach number of 0.8; in addition, the high- and low-speed engine operating limits were investigated up to the maximum operable altitude. The effect of changing the compressor bleed position on the stall and operating limits was determined for one of the inlet distortions. The circumferential distortions lowered the compressor stall pressure ratios; this resulted in less fuel-flow margin between steady-state operation and compressor stall. Consequently, the altitude operating Limits with circumferential distortions were reduced compared with the uniform inlet profile. Radial inlet-pressure distortions increased the pressure ratio required for compressor stall over that obtained with uniform inlet flow; this resulted in higher altitude operating limits. Likewise, the stall-limit fuel flows required with the radial inlet-pressure distortions were considerably higher than those obtained with the uniform inlet-pressure profile. A combined radial-circumferential inlet distortion had effects on the engine similar to the circumferential distortion. Bleeding air between the two compressors eliminated the low-speed stall limit and thus permitted higher altitude operation than was possible without compressor bleed.

Calculation of external-internal flow fields for mixed-compression inlets

NASA Technical Reports Server (NTRS)

Chyu, W. J.; Kawamura, T.; Bencze, D. P.

1986-01-01

Supersonic inlet flows with mixed external-internal compressions were computed using a combined implicit-explicit (Beam-Warming-Steger/MacCormack) method for solving the three-dimensional unsteady, compressible Navier-Stokes equations in conservation form. Numerical calculations were made of various flows related to such inlet operations as the shock-wave intersections, subsonic spillage around the cowl lip, and inlet started versus unstarted conditions. Some of the computed results were compared with wind tunnel data.

Calculation of external-internal flow fields for mixed-compression inlets

NASA Technical Reports Server (NTRS)

Chyu, W. J.; Kawamura, T.; Bencze, D. P.

1987-01-01

Supersonic inlet flows with mixed external-internal compressions were computed using a combined implicit-explicit (Beam-Warming-Steger/MacCormack) method for solving the three-dimensional unsteady, compressible Navier-Stokes equations in conservation form. Numerical calculations were made of various flows related to such inlet operations as the shock-wave intersections, subsonic spillage around the cowl lip, and inlet started versus unstarted conditions. Some of the computed results were compared with wind tunnel data.

Swozzle based burner tube premixer including inlet air conditioner for low emissions combustion

DOEpatents

Tuthill, Richard Sterling; Bechtel, II, William Theodore; Benoit, Jeffrey Arthur; Black, Stephen Hugh; Bland, Robert James; DeLeonardo, Guy Wayne; Meyer, Stefan Martin; Taura, Joseph Charles; Battaglioli, John Luigi

2002-01-01

A burner for use in a combustion system of a heavy-duty industrial gas turbine includes a fuel/air premixer having an air inlet, a fuel inlet, and an annular mixing passage. The fuel/air premixer mixes fuel and air into a uniform mixture for injection into a combustor reaction zone. The burner also includes an inlet flow conditioner disposed at the air inlet of the fuel/air premixer for controlling a radial and circumferential distribution of incoming air. The pattern of perforations in the inlet flow conditioner is designed such that a uniform air flow distribution is produced at the swirler inlet annulus in both the radial and circumference directions. The premixer includes a swozzle assembly having a series of preferably air foil shaped turning vanes that impart swirl to the airflow entering via the inlet flow conditioner. Each air foil contains internal fuel flow passages that introduce natural gas fuel into the air stream via fuel metering holes that pass through the walls of the air foil shaped turning vanes. By injecting fuel in this manner, an aerodynamically clean flow field is maintained throughout the premixer. By injecting fuel via two separate passages, the fuel/air mixture strength distribution can be controlled in the radial direction to obtain optimum radial concentration profiles for control of emissions, lean blow outs, and combustion driven dynamic pressure activity as machine and combustor load are varied.

Flow structure, heat transfer and pressure drop in varying aspect ratio two-pass rectangular smooth channels

NASA Astrophysics Data System (ADS)

Siddique, Waseem; El-Gabry, Lamyaa; Shevchuk, Igor V.; Hushmandi, Narmin B.; Fransson, Torsten H.

2012-05-01

Two-pass channels are used for internal cooling in a number of engineering systems e.g., gas turbines. Fluid travelling through the curved path, experiences pressure and centrifugal forces, that result in pressure driven secondary motion. This motion helps in moving the cold high momentum fluid from the channel core to the side walls and plays a significant role in the heat transfer in the channel bend and outlet pass. The present study investigates using Computational Fluid Dynamics (CFD), the flow structure, heat transfer enhancement and pressure drop in a smooth channel with varying aspect ratio channel at different divider-to-tip wall distances. Numerical simulations are performed in two-pass smooth channel with aspect ratio Win/H = 1:3 at inlet pass and Wout/H = 1:1 at outlet pass for a variety of divider-to-tip wall distances. The results show that with a decrease in aspect ratio of inlet pass of the channel, pressure loss decreases. The divider-to-tip wall distance (Wel) not only influences the pressure drop, but also the heat transfer enhancement at the bend and outlet pass. With an increase in the divider-to-tip wall distance, the areas of enhanced heat transfer shifts from side walls of outlet pass towards the inlet pass. To compromise between heat transfer and pressure drop in the channel, Wel/H = 0.88 is found to be optimum for the channel under study.

Influence of leading edge bluntness on hypersonic flow in a generic internal-compression inlet

NASA Astrophysics Data System (ADS)

Borovoy, V.; Egorov, I.; Mosharov, V.; Radchenko, V.; Skuratov, A.; Struminskaya, I.

2015-06-01

Flow and heat transfer inside a generic inlet are investigated experimentally. The cross section of the inlet is rectangular. The inlet is installed on a flat plat at a significant distance from the leading edge. The experiments are performed in TsAGI wind tunnel UT-1M working in the Ludwieg tube mode at Mach number M∞ = 5 and Reynolds numbers (based on the plate length L = 320 mm) Re∞L = 23 · 106 and 13 · 106. Steady flow duration is 40 ms. Optical panoramic methods are used for investigation of flow outside and inside the inlet as well. For this purpose, the cowl and one of two compressing wedges are made of a transparent material. Heat flux distribution is measured by thin luminescent Temperature Sensitive Paint (TSP). Surface flow and shear stress visualization is performed by viscous oil containing luminophor particles. The investigation shows that at high contraction ratio of the inlet, an increase of plate or cowl bluntness to some critical value leads to sudden change of the flow structure.

Investigation of Unsteady Flow Interaction Between an Ultra-Compact Inlet and a Transonic Fan

NASA Technical Reports Server (NTRS)

Hah, Chunill; Rabe, Douglas; Scribben, Angie

2015-01-01

In the study presented, unsteady flow interaction between an ultra-compact inlet and a transonic fan stage is investigated. Future combat aircraft engines require ultra-compact inlet ducts as part of an integrated, advanced propulsion system to improve air vehicle capability and effectiveness to meet future mission needs. The main purpose of the current study is to advance the understanding of the flow interaction between a modern ultra-compact inlet and a transonic fan for future design applications. Many experimental/ analytical studies have been reported on the aerodynamics of compact inlets in aircraft engines. On the other hand, very few studies have been reported on the effects of flow distortion from these inlets on the performance of the following fan/compressor stages. The primary goal of the study presented is to investigate how flow interaction between an ultra-compact inlet and a transonic compressor influence the operating margin of the compressor. Both Unsteady Reynolds-averaged Navier-Stokes (URANS) and Large Eddy Simulation (LES) approaches are used to calculate the unsteady flow field, and the numerical results are used to study the flow interaction. The present study indicates that stall inception of the following compressor stage is affected directly based on how the distortion pattern evolves before it interacts with the fan/compressor face. For the present compressor, the stall initiates at the tip section with clean inlet flow and distortion pattern away from the casing itself seems to have limited impacts on the stall inception of the compressor. A counter-rotating swirl, which is generated due to flow separation inside the s-shaped compact duct, generates an increased flow angle near the blade tip. This increased flow angle near the rotor tip due to the secondary flow from the counter-rotating vortices is the primary reason for the reduced compressor stall margin.

Effect of micro-scale wind on the measurement of airborne pollen concentrations using volumetric methods on a building rooftop

NASA Astrophysics Data System (ADS)

Miki, Kenji; Kawashima, Shigeto; Fujita, Toshio; Nakamura, Kimihito; Clot, Bernard

2017-06-01

Evaluating airborne pollen concentrations is important for the understanding of the spatiotemporal dispersion of pollen grains. Using two identical pollen monitors in parallel, we performed two experiments in order to study the influences of a) the physical characteristics (orientation) of the air inlet and b) the presence of obstacles in proximity to the monitors on airborne pollen concentration data. The first experiment consisted of an evaluation of airborne pollen concentrations using two different types of orifices; 1) a vertically oriented inlet and 2) a wind vane intake, both attached to the same type of automatic pollen sampler. The second experiment investigated the relationship between vertical wind speed and horizontal wind direction around an obstacle with the goal of studying the impact of micro-scale wind on pollen sampling efficiency. The results of the two experiments suggest that the wind path near an obstacle might be redirected in a vertical direction before or after the wind flows over the obstacle, which causes measurement errors of airborne pollen concentrations that are proportional to the vertical wind speed, especially when a vertically oriented inlet is used.

The effect of inlet boundary layer thickness on the flow within an annular S-shaped duct

DOE Office of Scientific and Technical Information (OSTI.GOV)

Sonoda, T.; Arima, T.; Oana, M.

1999-07-01

Experimental and numerical investigations were carried out to gain a better understanding of the flow characteristics within an annular S-shaped duct, including the effect of the inlet boundary layer (IBL) on the flow. A duct with six struts and the geometry as that used to connect compressor spools on the experimental small two-spool turbofan engine was investigated. A curved downstream annular passage with similar meridional flow path geometry to that of the centrifugal compressor has been fitted at the exit of S-shaped duct. Two types of the IBL (i.e., thin and thick IBL) were used. Results showed that large differencesmore » of flow patterns were observed at the S-shaped duct exit between two types of IBL, though the value of net total pressure loss has not been remarkably changed. According to overall total pressure loss, which includes the IBL loss, the total pressure loss was greatly increased near the hub as compared to that for a thin one. For the thick IBL, a vortex pair related to the hub-side horseshoe vortex and the separated flow found at the strut trailing edge has been clearly captured in the form of the total pressure loss contours and secondary flow vectors, experimentally and numerically. The high-pressure loss regions on either side of the strut wake near the hub may act on a downstream compressor performance. There is a much-distorted three-dimensional flow patterns at the exit of S-shaped duct. This means that the aerodynamic sensitivity of S-shaped duct to the IBL thickness is very high. Therefore, sufficient care is needed to design not only downstream aerodynamic components (for example, centrifugal impeller) but also upstream aerodynamic components (LPC OGV).« less

Euler Calculations at Off-Design Conditions for an Inlet of Inward Turning RBCC-SSTO Vehicle

NASA Technical Reports Server (NTRS)

Takashima, N.; Kothari, A. P.

1998-01-01

The inviscid performance of an inward turning inlet design is calculated computationally for the first time. Hypersonic vehicle designs based on the inward turning inlets have been shown analytically to have increased effective specific impulse and lower heat load than comparably designed vehicles with two-dimensional inlets. The inward turning inlets are designed inversely from inviscid stream surfaces of known flow fields. The computational study is performed on a Mach 12 inlet design to validate the performance predicted by the design code (HAVDAC) and calculate its off-design Mach number performance. The three-dimensional Euler equations are solved for Mach 4, 8, and 12 using a software package called SAM, which consists of an unstructured mesh generator (SAMmesh), a three-dimensional unstructured mesh flow solver (SAMcfd), and a CAD-based software (SAMcad). The computed momentum averaged inlet throat pressure is within 6% of the design inlet throat pressure. The mass-flux at the inlet throat is also within 7 % of the value predicted by the design code thereby validating the accuracy of the design code. The off-design Mach number results show that flow spillage is minimal, and the variation in the mass capture ratio with Mach number is comparable to an ideal 2-D inlet. The results from the inviscid flow calculations of a Mach 12 inward turning inlet indicate that the inlet design has very good on and off-design performance which makes it a promising design candidate for future air-breathing hypersonic vehicles.

Effects of seawater flow rate and evaporation temperature on performance of Sherbet type ice making machine

NASA Astrophysics Data System (ADS)

Son, C. H.; Yoon, J. I.; Choi, K. H.; Lee, H. K.; Lee, K. S.; Moon, C. G.; Seol, S. H.

2018-01-01

This study analyzes performance of the sherbet type ice making machine using seawater with respect to seawater volumetric flow rate, evaporation temperature, cooling water inlet and seawater inlet temperature as variables. Cooling water inlet and seawater inlet temperature are set considering average temperature of South Korea and the equator regions. Volumetric flow rate of seawater range is 0.75-1.75 LPM in this experiment. The results obtained from the experiment are as follows. As the seawater volumetric flow rate increases, or seawater inlet temperature increases, evaporation capacity tends to increase. At the point of seawater inlet temperature of 27°C and volumetric flow rate of 1.0LPM, evaporation capacity is over 2kW. On the other hand, results of COP change tendency are different from that of evaporation capacity. It appears to increase until volumetric flow rate of 1.0LPM, and decrease gradually from volumetric flow rate of 1.5LPM. This is due to the increase of compressor work to keep the evaporation pressure in accordance with the temperature of heat source. As the evaporation temperature decreases from -8 to -15°C, the evaporation capacity increases, but the COP decreases.

Experimental investigation of the draft tube inlet flow of a bulb turbine

NASA Astrophysics Data System (ADS)

Vuillemard, J.; Aeschlimann, V.; Fraser, R.; Lemay, S.; Deschênes, C.

2014-03-01

In the BulbT project framework, a bulb turbine model was studied with a strongly diverging draft tube. At high discharge, flow separation occurs in the draft tube correlated to significant efficiency and power drops. In this context, a focus was put on the draft tube inlet flow conditions. Actually, a precise inlet flow velocity field is required for comparison and validation purposes with CFD simulation. This paper presents different laser Doppler velocimetry (LDV) measurements at the draft tube inlet and their analysis. The LDV was setup to measure the axial and circumferential velocity on a radius under the runner and a diameter under the hub. A method was developed to perform indirect measurement of the mean radial velocity component. Five operating conditions were studied to correlate the inlet flow to the separation in the draft tube. Mean velocities, fluctuations and frequencies allowed characterizing the flow. Using this experimental database, the flow structure was characterized. Phase averaged velocities based on the runner position allowed detecting the runner blade wakes. The velocity gradients induced by the blade tip vortices were captured. The guide vane wakes was also detected at the draft tube inlet. The recirculation in the hub wake was observed.

Application of rotor mounted pressure transducers to analysis of inlet turbulence. [flow distortion in turbofan engine inlet

NASA Technical Reports Server (NTRS)

Hanson, D. B.

1976-01-01

Miniature pressure transducers installed near the leading edge of a fan blade were used to diagnose the non-uniform flow entering a subsonic tip speed turbofan on a static test stand. The pressure response of the blade to the inlet flow variations was plotted in a form which shows the space-time history of disturbances ingested by the rotor. Also, periodically sampled data values were auto- and cross-correlated as if they had been acquired from fixed hot wire anemometers at 150 equally spaced angles around the inlet. With a clean inlet and low wind, evidence of long, narrow turbulence eddies was easily found both in the boundary layer of the fan duct and outside the boundary layer. The role of the boundary layer was to follow and amplify disturbances in the outer flow. These eddies frequently moved around the inlet with a corkscrew motion as they passed through.

A pressure flux-split technique for computation of inlet flow behavior

NASA Technical Reports Server (NTRS)

Pordal, H. S.; Khosla, P. K.; Rubin, S. G.

1991-01-01

A method for calculating the flow field in aircraft engine inlets is presented. The phenomena of inlet unstart and restart are investigated. Solutions of the reduced Navier-Stokes (RNS) equations are obtained with a time consistent direct sparse matrix solver that computes the transient flow field both internal and external to the inlet. Time varying shocks and time varying recirculation regions can be efficiently analyzed. The code is quite general and is suitable for the computation of flow for a wide variety of geometries and over a wide range of Mach and Reynolds numbers.

A Computational and Experimental Investigation of a Three-Dimensional Hypersonic Scramjet Inlet Flow Field. Ph.D. Thesis

NASA Technical Reports Server (NTRS)

Holland, Scott Douglas

1991-01-01

A combined computational and experimental parametric study of the internal aerodynamics of a generic three dimensional sidewall compression scramjet inlet configuration was performed. The study was designed to demonstrate the utility of computational fluid dynamics as a design tool in hypersonic inlet flow fields, to provide a detailed account of the nature and structure of the internal flow interactions, and to provide a comprehensive surface property and flow field database to determine the effects of contraction ratio, cowl position, and Reynolds number on the performance of a hypersonic scramjet inlet configuration.

Pressure Recovery, Drag, and Subcritical Stability Characteristics of Three Conical Supersonic Diffusers at Stream Mach Numbers from 1.7 to 2.0

NASA Technical Reports Server (NTRS)

Nussdorfer, Theodore J; Obery, Leonard J; Englert, Gerald W

1952-01-01

A study of a 20 degree and a 25 degree half-angle high mass-flow ratio conical supersonic inlet was made on a 16-inch ram jet in the 8- by 6-foot supersonic tunnel. A greater range of stable subcritical operation was obtained with the low mass-flow ratio inlets; a greater range was obtained with the 25 degree than with the 20 degree half-angle low mass-flow ratio inlet. The high mass-flow ratio inlet had the least drag.

Light absorption cell combining variable path and length pump

DOEpatents

Prather, William S.

1993-01-01

A device for use in making spectrophotometric measurements of fluid samples. In particular, the device is a measurement cell containing a movable and a fixed lens with a sample of the fluid therebetween and through which light shines. The cell is connected to a source of light and a spectrophotometer via optic fibers. Movement of the lens varies the path length and also pumps the fluid into and out of the cell. Unidirectional inlet and exit valves cooperate with the movable lens to assure a one-way flow of fluid through the cell. A linear stepper motor controls the movement of the lens and cycles it from a first position closer to the fixed lens and a second position farther from the fixed lens, preferably at least 10 times per minute for a nearly continuous stream of absorption spectrum data.

Computational Modeling and Validation for Hypersonic Inlets

NASA Technical Reports Server (NTRS)

Povinelli, Louis A.

1996-01-01

Hypersonic inlet research activity at NASA is reviewed. The basis for the paper is the experimental tests performed with three inlets: the NASA Lewis Research Center Mach 5, the McDonnell Douglas Mach 12, and the NASA Langley Mach 18. Both three-dimensional PNS and NS codes have been used to compute the flow within the three inlets. Modeling assumptions in the codes involve the turbulence model, the nature of the boundary layer, shock wave-boundary layer interaction, and the flow spilled to the outside of the inlet. Use of the codes and the experimental data are helping to develop a clearer understanding of the inlet flow physics and to focus on the modeling improvements required in order to arrive at validated codes.

Effect of Inlet and Outlet Flow Conditions on Natural Gas Parameters in Supersonic Separation Process

PubMed Central

Yang, Yan; Wen, Chuang; Wang, Shuli; Feng, Yuqing

2014-01-01

A supersonic separator has been introduced to remove water vapour from natural gas. The mechanisms of the upstream and downstream influences are not well understood for various flow conditions from the wellhead and the back pipelines. We used a computational model to investigate the effect of the inlet and outlet flow conditions on the supersonic separation process. We found that the shock wave was sensitive to the inlet or back pressure compared to the inlet temperature. The shock position shifted forward with a higher inlet or back pressure. It indicated that an increasing inlet pressure declined the pressure recovery capacity. Furthermore, the shock wave moved out of the diffuser when the ratio of the back pressure to the inlet one was greater than 0.75, in which the state of the low pressure and temperature was destroyed, resulting in the re-evaporation of the condensed liquids. Natural gas would be the subsonic flows in the whole supersonic separator, if the mass flow rate was less than the design value, and it could not reach the low pressure and temperature for the condensation and separation of the water vapor. These results suggested a guidance mechanism for natural gas supersonic separation in various flow conditions. PMID:25338207

Sea level static calibration of a compact multimission aircraft propulsion simulator with inlet flow distortion

NASA Technical Reports Server (NTRS)

Won, Mark J.

1990-01-01

Wind tunnel tests of propulsion-integrated aircraft models have identified inlet flow distortion as a major source of compressor airflow measurement error in turbine-powered propulsion simulators. Consequently, two Compact Multimission Aircraft Propulsion Simulator (CMAPS) units were statically tested at sea level ambient conditions to establish simulator operating performance characteristics and to calibrate the compressor airflow against an accurate bellmouth flowmeter in the presence of inlet flow distortions. The distortions were generated using various-shaped wire mesh screens placed upstream of the compressor. CMAPS operating maps and performance envelopes were obtained for inlet total pressure distortions (ratio of the difference between the maximum and minimum total pressures to the average total pressure) up to 35 percent, and were compared to baseline simulator operating characteristics for a uniform inlet. Deviations from CMAPS baseline performance were attributed to the coupled variation of both compressor inlet-flow distortion and Reynolds number index throughout the simulator operating envelope for each screen configuration. Four independent methods were used to determine CMAPS compressor airflow; direct compressor inlet and discharge measurements, an entering/exiting flow-balance relationships, and a correlation between the mixer pressure and the corrected compressor airflow. Of the four methods, the last yielded the least scatter in the compressor flow coefficient, approximately + or - 3 percent over the range of flow distortions.

New design of a cathode flow-field with a sub-channel to improve the polymer electrolyte membrane fuel cell performance

NASA Astrophysics Data System (ADS)

Wang, Yulin; Yue, Like; Wang, Shixue

2017-03-01

The cathode flow-field design of polymer electrolyte membrane (PEM) fuel cells determines the distribution of reactant gases and the removal of liquid water. A suitable design can result in perfect water management and thus high cell performance. In this paper, a new design for a cathode flow-field with a sub-channel was proposed and had been experimentally analyzed in a parallel flow-field PEM fuel cell. Three sub-channel inlets were placed along the cathode channel. The main-channel inlet was fed with moist air to humidify the membrane and maintain high proton conductivity, whereas, the sub-channel inlet was fed with dry air to enhance water removal in the flow channel. The experimental results indicated that the sub-channel design can decrease the pressure drop in the flow channel, and the sub-channels inlet positions (SIP, where the sub-channel inlets were placed along the cathode channel) and flow rates (SFR, percentage of air from the sub-channel inlet in the total cathode flow rate) had a considerable impact on water removal and cell performance. A proposed design that combines the SIP and SFR can effectively eliminate water from the fuel cell, increasing the maximum power density by more than 13.2% compared to the conventional design.

Avco Lycoming QCGAT program design cycle, demonstrated performance and emissions

NASA Technical Reports Server (NTRS)

Fogel, P.; Koschier, A.

1980-01-01

A high bypass ratio, twin spool turbofan engine of modular design which incorporates a front fan module driven by a modified LTS101 core engine was tested. The engine is housed in a nacelle incorporating full length fan ducting with sound treatment in both the inlet and fan discharge flow paths. Design goals of components and results of component tests are presented together with full engine test results. The rationale behind the combustor design selected for the engine is presented as well as the emissions test results. Total system (engine and nacelle) test results are included.

Aerodynamic performance of a fan stage utilizing Variable Inlet Guide Vanes (VIGVs) for thrust modulation. [subsonic V/STOL aircraft

NASA Technical Reports Server (NTRS)

Woollett, R. R.

1983-01-01

An experimental research program was conducted in the Lewis Research Center's 9x15-foot (2.74x4.57 m) low speed wind tunnel to evaluate the aerodynamic performance of an inlet and fan system with variable inlet guide vanes (VIGVs) for use on a subsonic V/STOL aircraft. At high VIGV blade angles (lower weight flow and thrust levels), the fan stage was stalled over a major portion of its radius. In spite of the stall, fan blade stresses only exceeded the limits at the most extreme flow conditions. It was found that inlet flow separation does not necessarily lead to poor inlet performance or adverse fan operating conditions. Generally speaking, separated inlet flow did not adversely affect the fan blade stress levels. There were some cases, however, at high VIGV angles and high inlet angles-of-attack where excessive blade stress levels were encountered. An evaluation term made up of the product of the distortion parameter, K alpha, the weight flow and the fan pressure ratio minus one, was found to correlate quite well with the observed blade stress results.

Aerodynamic performance of a fan stage utilizing variable inlet guide vanes (VIGV's) for thrust modulation. [subsonic V/STOL aircraft

NASA Technical Reports Server (NTRS)

Woollett, R. R.

1983-01-01

An experimental research program was conducted in the Lewis Research Center's 9 x 15-foot (2.74 x 4.57 m) low speed wind tunnel to evaluate the aerodynamic performance of an inlet and fan system with variable inlet guide vanes (VIGVs) for use on a subsonic V/STOL aircraft. At high VIGV blade angles (lower weight flow and thrust levels), the fan stage was stalled over a major portion of its radius. In spite of the stall, fan blade stresses only exceeded the limits at the most extreme flow conditions. It was found that inlet flow separation does not necessarily lead to poor inlet performance or adverse fan operating conditions. Generally speaking, separated inlet flow did not adversely affect the fan blade stress levels. There were some cases, however, at high VIGV angles and high inlet angles-of-attack where excessive blade stress levels were encountered. An evaluation term made up of the product of the distortion parameter, K alpha, the weight flow and the fan pressure ratio minus one, was found to correlate quite well with the observed blade stress results. Previously announced in STAR as N83-27957

Users manual for updated computer code for axial-flow compressor conceptual design

NASA Technical Reports Server (NTRS)

Glassman, Arthur J.

1992-01-01

An existing computer code that determines the flow path for an axial-flow compressor either for a given number of stages or for a given overall pressure ratio was modified for use in air-breathing engine conceptual design studies. This code uses a rapid approximate design methodology that is based on isentropic simple radial equilibrium. Calculations are performed at constant-span-fraction locations from tip to hub. Energy addition per stage is controlled by specifying the maximum allowable values for several aerodynamic design parameters. New modeling was introduced to the code to overcome perceived limitations. Specific changes included variable rather than constant tip radius, flow path inclination added to the continuity equation, input of mass flow rate directly rather than indirectly as inlet axial velocity, solution for the exact value of overall pressure ratio rather than for any value that met or exceeded it, and internal computation of efficiency rather than the use of input values. The modified code was shown to be capable of computing efficiencies that are compatible with those of five multistage compressors and one fan that were tested experimentally. This report serves as a users manual for the revised code, Compressor Spanline Analysis (CSPAN). The modeling modifications, including two internal loss correlations, are presented. Program input and output are described. A sample case for a multistage compressor is included.

Fuel cell membrane hydration and fluid metering

DOEpatents

Jones, Daniel O.; Walsh, Michael M.

2003-01-01

A hydration system includes fuel cell fluid flow plate(s) and injection port(s). Each plate has flow channel(s) with respective inlet(s) for receiving respective portion(s) of a given stream of reactant fluid for a fuel cell. Each injection port injects a portion of liquid water directly into its respective flow channel. This serves to hydrate at least corresponding part(s) of a given membrane of the corresponding fuel cell(s). The hydration system may be augmented by a metering system including flow regulator(s). Each flow regulator meters an injecting at inlet(s) of each plate of respective portions of liquid into respective portion(s) of a given stream of fluid by corresponding injection port(s).

A study on flow development in an APU-style inlet and its effect on centrifugal compressor performance

NASA Astrophysics Data System (ADS)

Lou, Fangyuan

The objectives of this research were to investigate the flow development inside an APU-style inlet and its effect on centrifugal compressor performance. The motivation arises from the increased applications of gas turbine engines installed with APU-style inlets such as unmanned aerial vehicles, auxiliary power units, and helicopters. The inlet swirl distortion created from these complicated inlet systems has become a major performance and operability concern. To improve the integration between the APU-style inlet and gas turbine engines, better understanding of the flow field in the APU-style inlet and its effect on gas turbine is necessary. A research facility for the purpose of performing an experimental investigation of the flow field inside an APU-style inlet was developed. A subcritical air ejector is used to continuously flow the inlet at desired corrected mass flow rates. The facility is capable of flowing the APU inlet over a wide range of corrected mass flow rate that matches the same Mach numbers as engine operating conditions. Additionally, improvement in the system operational steadiness was achieved by tuning the pressure controller using a PID control method and utilizing multi-layer screens downstream of the APU inlet. Less than 1% relative unsteadiness was achieved for full range operation. The flow field inside the rectangular-sectioned 90? bend of the APU-style inlet was measured using a 3-Component LDV system. The structures for both primary flow and the secondary flow inside the bend were resolved. Additionally, the effect of upstream geometry on the flow development in the downstream bend was also investigated. Furthermore, a Single Stage Centrifugal Compressor research facility was developed at Purdue University in collaboration with Honeywell to operate the APU-style inlet at engine conditions with a compressor. To operate the facility, extensive infrastructure for facility health monitoring and performance control (including lubrication systems, secondary air systems, a throttle system, and different inlet configurations) were built. Additionally, three Labview programs were developed for acquiring the compressor health monitoring, steady and unsteady pressure and strain data. The baseline, steady aerodynamic performance map was established. Additionally, the unsteady pressure field in the compressor was investigated. Steady performance data have been acquired from choke to near surge at three different corrected speeds from 90% to 100% corrected speed in 5% increments. The performance of the compressor stage was characterized using total pressure ratio (TPR), total temperature ratio (TTR), and isentropic efficiency. The impeller alone and diffuser along performance were also investigated, and the high loss regions in the compressor were identified. At last, the compressor unsteady shroud pressure was investigated at 100% corrected speed in both the time domain and frequency domain. Results show strong pressure components in relation to the shaft frequency (SF). The impeller has 17 main blades and 17 splitter blades, and introduces pressure fluctuations at 17SF and its harmonics. Additionally, the diffuser has a vane count of 25 and results in pressure spectra of 59SF (17+17+25) due to the interactions between the impeller and diffuser.

Analysis of the flow field generated near an aircraft engine operating in reverse thrust. M.S. Thesis

NASA Technical Reports Server (NTRS)

Ledwith, W. A., Jr.

1972-01-01

A computer solution is developed to the exhaust gas reingestion problem for aircraft operating in the reverse thrust mode on a crosswind-free runway. The computer program determines the location of the inlet flow pattern, whether the exhaust efflux lies within the inlet flow pattern or not, and if so, the approximate time before the reversed flow reaches the engine inlet. The program is written so that the user is free to select discrete runway speeds or to study the entire aircraft deceleration process for both the far field and cross-ingestion problems. While developed with STOL applications in mind, the solution is equally applicable to conventional designs. The inlet and reversed jet flow fields involved in the problem are assumed to be noninteracting. The nacelle model used in determining the inlet flow field is generated using an iterative solution to the Neuman problem from potential flow theory while the reversed jet flow field is adapted using an empirical correlation from the literature. Sample results obtained using the program are included.

The 3-D viscous flow CFD analysis of the propeller effect on an advanced ducted propeller subsonic inlet

NASA Technical Reports Server (NTRS)

Iek, Chanthy; Boldman, Donald R.; Ibrahim, Mounir

1993-01-01

A time marching Navier-Stokes code called PARC3D was used to study the 3-D viscous flow associated with an advanced ducted propeller (ADP) subsonic inlet at take-off operating conditions. At a free stream Mach number of 0.2, experimental data for the inlet-with-propeller test model indicated that the airflow was attached on the cowl windward lip at an angle of attack of 25 degrees became unstable at 29 degrees, and separated at 30 degrees. An experimental study with a similar inlet and with no propeller (through-flow) indicated that flow separation occurred at an angle of attack a few degrees below the value observed when the inlet was tested with the propeller. This tends to indicate that the propeller exerts a favorable effect on the inlet performance. During the through-flow experiment a stationary blockage device was used to successfully simulate the propeller effect on the inlet flow field at angles of attack. In the present numerical study, this flow blockage was modeled via a PARC3D computational boundary condition (BC) called the screen BC. The principle formulation of this BC was based on the one-and-half dimension actuator disk theory. This screen BC was applied at the inlet propeller face station of the computational grid. Numerical results were obtained with and without the screen BC. The application of the screen BC in this numerical study provided results which are similar to the results of past experimental efforts in which either the blockage device or the propeller was used.

Active Flow Control on a Boundary-Layer-Ingesting Inlet

NASA Technical Reports Server (NTRS)

Gorton, Susan Althoff; Owens, Lewis R.; Jenkins, Luther N.; Allan, Brian G.; Schuster, Ernest P.

2004-01-01

Boundary layer ingestion (BLI) is explored as means to improve overall system performance for Blended Wing Body configuration. The benefits of BLI for vehicle system performance benefit are assessed with a process derived from first principles suitable for highly-integrated propulsion systems. This performance evaluation process provides framework within which to assess the benefits of an integrated BLI inlet and lays the groundwork for higher-fidelity systems studies. The results of the system study show that BLI provides a significant improvement in vehicle performance if the inlet distortion can be controlled, thus encouraging the pursuit of active flow control (AFC) as a BLI enabling technology. The effectiveness of active flow control in reducing engine inlet distortion was assessed using a 6% scale model of a 30% BLI offset, diffusing inlet. The experiment was conducted in the NASA Langley Basic Aerodynamics Research Tunnel with a model inlet designed specifically for this type of testing. High mass flow pulsing actuators provided the active flow control. Measurements were made of the onset boundary layer, the duct surface static pressures, and the mass flow through the duct and the actuators. The distortion was determined by 120 total pressure measurements located at the aerodynamic interface plane. The test matrix was limited to a maximum freestream Mach number of 0.15 with scaled mass flows through the inlet for that condition. The data show that the pulsed actuation can reduce distortion from 29% to 4.6% as measured by the circumferential distortion descriptor DC60 using less than 1% of inlet mass flow. Closed loop control of the actuation was also demonstrated using a sidewall surface static pressure as the response sensor.

Aerothermodynamic flow phenomena of the airframe-integrated supersonic combustion ramjet

NASA Technical Reports Server (NTRS)

Walton, James T.

1992-01-01

The unique component flow phenomena is discussed of the airframe-integrated supersonic combustion ramjet (scramjet) in a format geared towards new players in the arena of hypersonic propulsion. After giving an overview of the scramjet aerothermodynamic cycle, the characteristics are then covered individually of the vehicle forebody, inlet, combustor, and vehicle afterbody/nozzle. Attention is given to phenomena such as inlet speeding, inlet starting, inlet spillage, fuel injection, thermal choking, and combustor-inlet interaction.

Flow Simulation of Supersonic Inlet with Bypass Annular Duct

NASA Technical Reports Server (NTRS)

Kim, HyoungJin; Kumano, Takayasu; Liou, Meng-Sing; Povinelli, Louis A.; Conners, Timothy R.

2011-01-01

A relaxed isentropic compression supersonic inlet is a new concept that produces smaller cowl drag than a conventional inlet, but incurs lower total pressure recovery and increased flow distortion in the (radially) outer flowpath. A supersonic inlet comprising a bypass annulus to the relaxed isentropic compression inlet dumps out airflow of low quality through the bypass duct. A reliable computational fluid dynamics solution can provide considerable useful information to ascertain quantitatively relative merits of the concept, and further provide a basis for optimizing the design. For a fast and reliable performance evaluation of the inlet performance, an equivalent axisymmetric model whose area changes accounts for geometric and physical (blockage) effects resulting from the original complex three-dimensional configuration is proposed. In addition, full three-dimensional calculations are conducted for studying flow phenomena and verifying the validity of the equivalent model. The inlet-engine coupling is carried out by embedding numerical propulsion system simulation engine data into the flow solver for interactive boundary conditions at the engine fan face and exhaust plane. It was found that the blockage resulting from complex three-dimensional geometries in the bypass duct causes significant degradation of inlet performance by pushing the terminal normal shock upstream.

Several examples where turbulence models fail in inlet flow field analysis

NASA Technical Reports Server (NTRS)

Anderson, Bernhard H.

1993-01-01

Computational uncertainties in turbulence modeling for three dimensional inlet flow fields include flows approaching separation, strength of secondary flow field, three dimensional flow predictions of vortex liftoff, and influence of vortex-boundary layer interactions; computational uncertainties in vortex generator modeling include representation of generator vorticity field and the relationship between generator and vorticity field. The objectives of the inlet flow field studies presented in this document are to advance the understanding, prediction, and control of intake distortion and to study the basic interactions that influence this design problem.

Effect of self recirculation casing treatment on the performance of a turbocharger centrifugal compressor

NASA Astrophysics Data System (ADS)

Gancedo, Matthieu

Increase in emission regulations in the transport industry brings the need to have more efficient engines. A path followed by the automobile industry is to downsize the size of the internal combustion engine and increase the air density at the intake to keep the engine power when needed. Typically a centrifugal compressor is used to force the air into the engine, it can be powered from the engine shaft (superchargers) or extracting energy contained into the hot exhaust gases with a turbine (turbochargers). The flow range of the compressor needs to match the one of the engine. However compressors mass flow operating range is limited by choke on the high end and surge on the low end. In order to extend the operation at low mass flow rates, the use of passive devices for turbocharger centrifugal compressors was explored since the late 80's. Hence, casing treatments including flow recirculation from the inducer part of the compressor have been shown to move the surge limit to lower flows. Yet, the working mechanisms are still not well understood and thus, to optimize the design of this by-pass system, it is necessary to determine the nature of the changes induced by the device both on the dynamic stability of the pressure delivery and on the flow at the inlet. The compressor studied here features a self-recirculating casing treatment at the inlet. The recirculation passage could be blocked to carry a direct comparison between the cases with and without the flow feature. To grasp the effect on compressor stability, pressure measurements were taken in the different constituting elements of the compressor. The study of the mean pressure variations across the operating map showed that the tongue region is a limiting element. Dynamic pressure measurements revealed that the instabilities generated near the inducer when the recirculation is blocked increase the overall instability levels at the compressor outlet and propagating pressure waves starting at the tongue occurred, different in nature from rotating stall. The flow velocity was also measured at the inlet of the compressor by means of planar PIV measurements. The case without recirculation showed strong back flow occurrence at low MFR on the shroud of the inlet passage due to tip recirculation. With recirculation, this back flow was significantly reduced improving the overall stability. However, with the current recirculation channels design, there is an efficiency penalty and the recirculated flow introduces non-homogeneities in the mixing region. Finally, to explore experimentally the effect of variations of the casing treatment, several different designs were tested. It was seen that modifications of the supporting rib shape impacted the efficiency. Also, improvements on the surge line were obtained with flow reinjection near the inducer in the direction of the main flow at low speeds and with induced counter swirl for all speeds.

Unstart Coupling Mechanism Analysis of Multiple-Modules Hypersonic Inlet

PubMed Central

Wang, Lei; Cao, Shibin

2013-01-01

The combination of multiplemodules in parallel manner is an important way to achieve the much higher thrust of scramjet engine. For the multiple-modules scramjet engine, when inlet unstarted oscillatory flow appears in a single-module engine due to high backpressure, how to interact with each module by massflow spillage, and whether inlet unstart occurs in other modules are important issues. The unstarted flowfield and coupling characteristic for a three-module hypersonic inlet caused by center module II and side module III were, conducted respectively. The results indicate that the other two hypersonic inlets are forced into unstarted flow when unstarted phenomenon appears on a single-module hypersonic inlet due to high backpressure, and the reversed flow in the isolator dominates the formation, expansion, shrinkage, and disappearance of the vortexes, and thus, it is the major factor of unstart coupling of multiple-modules hypersonic inlet. The coupling effect among multiple modules makes hypersonic inlet be more likely unstarted. PMID:24348146

Time-lapse 3D imaging of calcite precipitation in a microporous column

NASA Astrophysics Data System (ADS)

Godinho, Jose R. A.; Withers, Philip J.

2018-02-01

Time-lapse X-ray computed tomography is used to image the evolution of calcite precipitation during flow through microporous quartz over the course of 400 h. The growth rate decreases by more than seven times, which is linked to the clogging of flow paths that restricts flow to some regions of the column. Fewer precipitates are observed as a function of column depth, which is found to be related to a differential nucleation density along the sample. A higher nucleation density closer to the inlet implies more crystal volume increase per unit of time without affecting the rate if normalized to the surface area of crystals. Our overall growth rates measured in porous media are orders of magnitude slower than growth rates derived from traditional precipitation experiments on free surfaces. Based on our time-lapse results we hypothesize a scenario where the evolving distribution of precipitates within a pore structure during precipitation progressively modifies the local transport through the pores. Within less permeable regions the saturation index may be lower than along the main flow paths. Therefore, the reactive crystal surfaces within those regions grow at a slower rate than that expected from the bulk fluid composition. Since the amount of reactive surface area within these less permeable regions increases over time, the overall growth rate decreases without a necessary significant change of the bulk fluid composition along more permeable flow paths. In conclusion, the overall growth rates in an evolving porous media expected from bulk fluid compositions alone can be overestimated due to the development of stagnant sub-regions where the reactive surface area is bath by a solution with lower saturation index. In this context we highlight the value of time-lapse 3D studies for understanding the dynamics of mineral precipitation in porous media.

Investigation of Unsteady Flow Interaction Between an Ultra-Compact Inlet and a Transonic Fan

NASA Technical Reports Server (NTRS)

Hah, Chunill; Rabe, Douglas; Scribben, Angie

2015-01-01

In the present study, unsteady flow interaction between an ultra-compact inlet and a transonic fan stage is investigated. Future combat aircraft require ultra-compact inlet ducts as part of an integrated, advanced propulsion system to improve air vehicle capability and effectiveness to meet future mission needs. The main purpose of the study is to advance the current understanding of the flow interaction between two different ultra-compact inlets and a transonic fan for future design applications. Both URANS and LES approaches are used to calculate the unsteady flow field and are compared with the available measured data. The present study indicates that stall inception is mildly affected by the distortion pattern generated by the inlet with the current test set-up. The numerical study indicates that the inlet distortion pattern decays significantly before it reaches the fan face for the current configuration. Numerical results with a shorter distance between the inlet and fan show that counter-rotating vortices near the rotor tip due to the serpentine diffuser affects fan characteristics significantly.

Advanced two-stage compressor program design of inlet stage

NASA Technical Reports Server (NTRS)

Bryce, C. A.; Paine, C. J.; Mccutcheon, A. R. S.; Tu, R. K.; Perrone, G. L.

1973-01-01

The aerodynamic design of an inlet stage for a two-stage, 10/1 pressure ratio, 2 lb/sec flow rate compressor is discussed. Initially a performance comparison was conducted for an axial, mixed flow and centrifugal second stage. A modified mixed flow configuration with tandem rotors and tandem stators was selected for the inlet stage. The term conical flow compressor was coined to describe a particular type of mixed flow compressor configuration which utilizes axial flow type blading and an increase in radius to increase the work input potential. Design details of the conical flow compressor are described.

Orifice Mass Flow Calculation in NASA's W-8 Single Stage Axial Compressor Facility

NASA Technical Reports Server (NTRS)

Bozak, Richard F.

2018-01-01

Updates to the orifice mass flow calculation for the W-8 Single Stage Axial Compressor Facility at NASA Glenn Research Center are provided to include the effect of humidity and incorporate ISO 5167. A methodology for including the effect of humidity into the inlet orifice mass flow calculation is provided. Orifice mass flow calculations provided by ASME PTC-19.5-2004, ASME MFC-3M-2004, ASME Fluid Meters, and ISO 5167 are compared for W-8's atmospheric inlet orifice plate. Differences in expansion factor and discharge coefficient given by these standards give a variation of about +/- 75% mass flow except for a few cases. A comparison of the calculations with an inlet static pressure mass flow correlation and a fan exit mass flow integration using test data from a 2017 turbofan rotor test in W-8 show good agreement between the inlet static pressure mass flow correlation, ISO 5167, and ASME Fluid Meters. While W-8's atmospheric inlet orifice plate violates the pipe diameter limit defined by each of the standards, the ISO 5167 is chosen to be the primary orifice mass flow calculation to use in the W-8 facility.

Complex fluid flow and heat transfer analysis inside a calandria based reactor using CFD technique

NASA Astrophysics Data System (ADS)

Kulkarni, P. S.

2017-04-01

Series of numerical experiments have been carried out on a calandria based reactor for optimizing the design to increase the overall heat transfer efficiency by using Computational Fluid Dynamic (CFD) technique. Fluid flow and heat transfer inside the calandria is governed by many geometric and flow parameters like orientation of inlet, inlet mass flow rate, fuel channel configuration (in-line, staggered, etc.,), location of inlet and outlet, etc.,. It was well established that heat transfer is more wherever forced convection dominates but for geometries like calandria it is very difficult to achieve forced convection flow everywhere, intern it strongly depends on the direction of inlet jet. In the present paper the initial design was optimized with respect to inlet jet angle, the optimized design has been numerically tested for different heat load mass flow conditions. To further increase the heat removal capacity of a calandria, further numerical studies has been carried out for different inlet geometry. In all the analysis same overall geometry size and same number of tubes has been considered. The work gives good insight into the fluid flow and heat transfer inside the calandria and offer a guideline for optimizing the design and/or capacity enhancement of a present design.

Control of reactor coolant flow path during reactor decay heat removal

DOEpatents

Hunsbedt, Anstein N.

1988-01-01

An improved reactor vessel auxiliary cooling system for a sodium cooled nuclear reactor is disclosed. The sodium cooled nuclear reactor is of the type having a reactor vessel liner separating the reactor hot pool on the upstream side of an intermediate heat exchanger and the reactor cold pool on the downstream side of the intermediate heat exchanger. The improvement includes a flow path across the reactor vessel liner flow gap which dissipates core heat across the reactor vessel and containment vessel responsive to a casualty including the loss of normal heat removal paths and associated shutdown of the main coolant liquid sodium pumps. In normal operation, the reactor vessel cold pool is inlet to the suction side of coolant liquid sodium pumps, these pumps being of the electromagnetic variety. The pumps discharge through the core into the reactor hot pool and then through an intermediate heat exchanger where the heat generated in the reactor core is discharged. Upon outlet from the heat exchanger, the sodium is returned to the reactor cold pool. The improvement includes placing a jet pump across the reactor vessel liner flow gap, pumping a small flow of liquid sodium from the lower pressure cold pool into the hot pool. The jet pump has a small high pressure driving stream diverted from the high pressure side of the reactor pumps. During normal operation, the jet pumps supplement the normal reactor pressure differential from the lower pressure cold pool to the hot pool. Upon the occurrence of a casualty involving loss of coolant pump pressure, and immediate cooling circuit is established by the back flow of sodium through the jet pumps from the reactor vessel hot pool to the reactor vessel cold pool. The cooling circuit includes flow into the reactor vessel liner flow gap immediate the reactor vessel wall and containment vessel where optimum and immediate discharge of residual reactor heat occurs.

Development of PZT Actuated Valveless Micropump.

PubMed

Munas, Fathima Rehana; Melroy, Gehan; Abeynayake, Chamitha Bhagya; Chathuranga, Hiniduma Liyanage; Amarasinghe, Ranjith; Kumarage, Pubudu; Dau, Van Thanh; Dao, Dzung Viet

2018-04-24

A piezoelectrically actuated valveless micropump has been designed and developed. The principle components of this system are piezoelectrically actuated (PZT) metal diaphragms and a complete fluid flow system. The design of this pump mainly focuses on a cross junction, which is generated by a nozzle jet attached to a pump chamber and the intersection of two inlet channels and an outlet channel respectively. During each PZT diaphragm vibration cycle, the junction connecting the inlet and outlet channels with the nozzle jet permits consistencies in fluidic momentum and resistances in order to facilitate complete fluidic path throughout the system, in the absence of any physical valves. The entire micropump structure is fabricated as a plate-by-plate element of polymethyl methacrylate (PMMA) sheets and sandwiched to get required fluidic network as well as the overall device. In order to identify the flow characteristics, and to validate the test results with numerical simulation data, FEM analysis using ANSYS was carried out and an eigenfrequency analysis was performed to the PZT diaphragm using COMSOL Multiphysics. In addition, the control system of the pump was designed and developed to change the applied frequency to the piezoelectric diaphragms. The experimental data revealed that the maximum flow rate is 31.15 mL/min at a frequency of 100 Hz. Our proposed design is not only for a specific application but also useful in a wide range of biomedical applications.

Single stage experimental evaluation of tandem-airfoil rotor and stator blading for compressors. Part 4: Data and performance for stage B

NASA Technical Reports Server (NTRS)

Brent, J. A.; Cheatham, J. G.

1973-01-01

Stage B, composed of tandem-airfoil rotor B and stator B, was tested with uniform inlet flow and with hub radial, tip radial and 90 degree one-per-revolution circumferential distortion of the inlet flow as part of an overall program to evaluate the effectiveness of tandem airfoils for increasing the design point loading capability and stable operating range of rotor and stator blading. The results of this series of tests provide overall performance and blade element data for evaluating: (1) the potential of tandem blading for extending the loading limit and stable operating range of a stage representative of a middle stage of an advanced high pressure compressor, (2) the effect of loading split between the two airfoils in tandem on the performance of tandem blading, and (3) the effects of inlet flow distortion on the stage performance. The rotor had an inlet hub/tip ratio of 0.8 and a design tip velocity of 757 ft/sec. With uniform inlet flow, rotor B achieved a maximum adiabatic efficiency of 88.4% at design equivalent rotor speed and a pressure ratio of 1.31. The stage maximum adiabatic efficiency at design equivalent rotor speed with uniform inlet flow was 82.5% at a pressure ratio of 1.28. Tip radial and circumferential distortion of the inlet flow caused substantial reductions in surge margin.

Investigation of modification design of the fan stage in axial compressor

NASA Astrophysics Data System (ADS)

Zhou, Xun; Yan, Peigang; Han, Wanjin

2010-04-01

The S2 flow path design method of the transonic compressor is used to design the one stage fan in order to replace the original designed blade cascade which has two-stage transonic fan rotors. In the modification design, the camber line is parameterized by a quartic polynomial curve and the thickness distribution of the blade profile is controlled by the double-thrice polynomial. Therefore, the inlet flow has been pre-compressed and the location and intensity of the shock wave at supersonic area have been controlled in order to let the new blade profiles have better aerodynamic performance. The computational results show that the new single stage fan rotor increases the efficiency by two percent at the design condition and the total pressure ratio is slightly higher than that of the original design. At the same time, it also meets the mass flow rate and the geometrical size requirements for the modification design.

Evaluation of Flush-Mounted, S-Duct Inlets With Large Amounts of Boundary Layer Ingestion

NASA Technical Reports Server (NTRS)

Berrier, Bobby L.; Morehouse, Melissa B.

2003-01-01

A new high Reynolds number test capability for boundary layer ingesting inlets has been developed for the NASA Langley Research Center 0.3-Meter Transonic Cryogenic Tunnel. Using this new capability, an experimental investigation of four S-duct inlet configurations with large amounts of boundary layer ingestion (nominal boundary layer thickness of about 40% of inlet height) was conducted at realistic operating conditions (high subsonic Mach numbers and full-scale Reynolds numbers). The objectives of this investigation were to 1) develop a new high Reynolds number, boundary-layer ingesting inlet test capability, 2) evaluate the performance of several boundary layer ingesting S-duct inlets, 3) provide a database for CFD tool validation, and 4) provide a baseline inlet for future inlet flow-control studies. Tests were conducted at Mach numbers from 0.25 to 0.83, Reynolds numbers (based on duct exit diameter) from 5.1 million to a fullscale value of 13.9 million, and inlet mass-flow ratios from 0.39 to 1.58 depending on Mach number. Results of this investigation indicate that inlet pressure recovery generally decreased and inlet distortion generally increased with increasing Mach number. Except at low Mach numbers, increasing inlet mass-flow increased pressure recovery and increased distortion. Increasing the amount of boundary layer ingestion (by decreasing inlet throat height and increasing inlet throat width) or ingesting a boundary layer with a distorted profile decreased pressure recovery and increased distortion. Finally, increasing Reynolds number had almost no effect on inlet distortion but increased inlet recovery by about one-half percent at a Mach number near cruise.

Performance of the University of Denver Low Turbulence, Airborne Aerosol Inlet in ACE-Asia

NASA Astrophysics Data System (ADS)

Lafleur, B.; Wilson, J. C.; Seebaugh, W. R.; Gesler, D.; Hilbert, H.; Mullen, J.; Reeves, J. M.

2002-12-01

The University of Denver Low Turbulence Inlet (DULTI) was flown on the NCAR C-130 in ACE-Asia. This inlet delivered large sample flows at velocities of a few meters per second at the exit of the inlet. This flow was slowed from the true air speed of the aircraft (100 to 150 m/s) to a few meters per second in a short diffuser with porous walls. The flow in the diffusing section was laminar. The automatic control system kept the inlet operating at near isokinetic intake velocities and in laminar flow for nearly all the flight time. The DULTI permits super micron particles to be sampled and delivered with high efficiency to the interior of the aircraft where they can be measured or collected. Because most of the air entering the inlet is removed through the porous medium, the sample flow experiences inertial enhancements. Because these enhancements occur in laminar flow, they are calculable using FLUENT. Enhancement factors are defined as the ratio of the number of particles of a given size per unit mass of air in the sample to the number of particles of that size per unit mass of air in the ambient. Experimenters divide measured mixing ratios of the aerosol by the enhancement factor to get the ambient mixing ratio of the particles. The diffuser used in ACE-Asia differed from that used in PELTI (2000), TexAQS2000 (2000) and ITCT (2002). In this poster, the flow parameters measured in the inlet in flight are compared with those calculated from FLUENT. And enhancement factors are presented for flight conditions. The enhancement factors are found to depend upon the Stokes number of particles in the entrance to the inlet and the ratio of the mass flow rate of air removed by suction to the mass flow rate delivered as sample.

Evaluation of Flush-Mounted, S-Duct Inlets with Large Amounts of Boundary Layer Ingestion

NASA Technical Reports Server (NTRS)

Berrier, Bobby L.; Morehouse, Melissa B.

2003-01-01

A new high Reynolds number test capability for boundary layer ingesting inlets has been developed for the NASA Langley Research Center 0.3-Meter Transonic Cryogenic Tunnel. Using this new capability, an experimental investigation of four S-duct inlet configurations with large amounts of boundary layer ingestion (nominal boundary layer thickness of about 40% of inlet height) was conducted at realistic operating conditions (high subsonic Mach numbers and full-scale Reynolds numbers). The objectives of this investigation were to 1) provide a database for CFD tool validation on boundary layer ingesting inlets operating at realistic conditions and 2) provide a baseline inlet for future inlet flow-control studies. Tests were conducted at Mach numbers from 0.25 to 0.83, Reynolds numbers (based on duct exit diameter) from 5.1 million to a full-scale value of 13.9 million, and inlet mass-flow ratios from 0.39 to 1.58 depending on Mach number. Results of this investigation indicate that inlet pressure recovery generally decreased and inlet distortion generally increased with increasing Mach number. Except at low Mach numbers, increasing inlet mass-flow increased pressure recovery and increased distortion. Increasing the amount of boundary layer ingestion (by decreasing inlet throat height) or ingesting a boundary layer with a distorted (adverse) profile decreased pressure recovery and increased distortion. Finally, increasing Reynolds number had almost no effect on inlet distortion but increased inlet recovery by about one-half percent at a Mach number near cruise.

Two-Dimensional Bifurcated Inlet Variable Cowl Lip Test Completed in 10- by 10-Foot Supersonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Hoffman, T. R.

2000-01-01

Researchers at the NASA Glenn Research Center at Lewis Field successfully tested a variable cowl lip inlet at simulated takeoff conditions in Glenn s 10- by 10-Foot Supersonic Wind Tunnel (10x10 SWT) as part of the High-Speed Research Program. The test was a follow-on to the Two-Dimensional Bifurcated (2DB) Inlet/Engine test. At the takeoff condition for a High-Speed Civil Transport aircraft, the inlet must provide adequate airflow to the engine with an acceptable distortion level and high-pressure recovery. The test was conducted to study the effectiveness of installing two rotating lips on the 2DB Inlet cowls to increase mass flow rate and eliminate or reduce boundary layer flow separation near the lips. Hardware was mounted vertically in the test section so that it extended through the tunnel ceiling and that the 2DB Inlet was exposed to the atmosphere above the test section. The tunnel was configured in the aerodynamic mode, and exhausters were used to pump down the tunnel to vacuum levels and to provide a maximum flow rate of approximately 58 lb/sec. The test determined the (1) maximum flow in the 2DB Inlet for each variable cowl lip, (2) distortion level and pressure recovery for each lip configuration, (3) boundary layer conditions near variable lips inside the 2DB Inlet, (4) effects of a wing structure adjacent to the 2DB Inlet, and (5) effects of different 2DB Inlet exit configurations. It also employed flow visualization to generate enough qualitative data on variable lips to optimize the variable lip concept. This test was a collaborative effort between the Boeing Company and Glenn. Extensive inhouse support at Glenn contributed significantly to the progress and accomplishment of this test.

Numerical investigations in three-dimensional internal flows

NASA Technical Reports Server (NTRS)

Rose, William C.

1991-01-01

In previous efforts, a two-dimensional full Navier-Stokes (FNS) code (SCRAM2D) was used in a design process that involved parametric modifications of the inlet geometry to arrive at what appeared to be an optimum inlet flowfield that produced a uniform flow at the exit in a very short distance. In these previous studies, the technologies for determining the contours with a 'man-in-the-loop' approach for both the ramp and cowl of the inlet were demonstrated, and nearly shock-free exiting flowfields were shown to be obtainable. The resulting two-dimensional compression contours were then used with swept sidewalls to form a three-dimensional inlet. Then the three-dimensional Navier-Stokes code (SCRAM3D) was used to investigate the inlet's three-dimensional flow. One of the major difficulties encountered in the previous studies was that associated with the relatively long time required to obtain a solution using even the 2D FNS code in the design process. Since one of the goals of high-speed inlet design is to produce inputs to the overall aircraft design in a timely manner, it was proposed for this year's research to examine 2D and 3D viscous flow solver techniques alternative to the NFS codes used to date. Areas of the inlet particularly identified for code speed up are those associated with the forebody and external flow ramp systems of the inlet. In these areas, parabolized, or space-marched, Navier-Stokes codes were proposed to be investigated for their applicability in the design process developed previously. This report describes the results of an investigation into the use of two other codes for analyzing the forebody and inlet ramp systems of high-speed inlets.

Flow visualization study of the HiMAT RPRV

NASA Technical Reports Server (NTRS)

Lorincz, D. J.

1980-01-01

Water tunnel studies were performed to qualitatively define the flow field of the highly maneuverable aircraft technology remotely piloted research vehicle (HiMAT RPRV). Particular emphasis was placed on defining the vortex flows generated at high angles of attack. The flow visualization tests were conducted in the Northrop water tunnel using a 1/15 scale model of the HiMAT RPRV. Flow visualization photographs were obtained for angles of attack up to 40 deg and sideslip angles up to 5 deg. The HiMAT model was investigated in detail to determine the canard and wing vortex flow field development, vortex paths, and vortex breakdown characteristics as a function of angle of attack and sideslip. The presence of the canard caused the wing vortex to form further outboard and delayed the breakdown of the wing vortex to higher angles of attack. An increase in leading edge camber of the maneuver configuration delayed both the formation and the breakdown of the wing and canard vortices. Additional tests showed that the canard vortex was sensitive to variations in inlet mass flow ratio and canard flap deflection angle.

Ducted fan inlet/exit and rotor tip flow improvements for vertical lift systems

NASA Astrophysics Data System (ADS)

Akturk, Ali

The current research utilized experimental and computational techniques in 5" and 22" diameter ducted fan test systems that have been custom designed and manufactured. Qualitative investigation of flow around the ducted fan was also performed using smoke flow visualizations. Quantitative measurements consisted of 2D and 3D velocity measurements using planar and Stereoscopic Particle Image Velocimetry (PIV and SPIV), high resolution total pressure measurements using Kiel total pressure probes and real time six-component force and torque measurements. The computational techniques used in this thesis included a recently developed radial equilibrium based rotor model(REBRM) and a three dimensional Reynolds-Averaged Navier Stokes (RANS) based CFD model. A radial equilibrium based rotor model (REBRM) developed by the author was effectively integrated into a three-dimensional RANS based computational system. The PIV measurements and computational flow predictions using (REBRM) near the fan inlet plane were in a good agreement at hover and forward flight conditions. The aerodynamic modifications resulting from the fan inlet flow distortions in forward flight regime were clearly captured in 2D PIV results. High resolution total pressure measurements at the downstream of the fan rotor showed that tip leakage, rotor hub separation, and passage flow related total pressure losses were dominant in hover condition. However, the losses were dramatically increased in forward flight because of inlet lip separation and distortion. A novel ducted fan inlet flow conditioning concept named "Double Ducted Fan" (DDF) was developed. The (DDF) concept has a potential to significantly improve the performance and controllability of VTOL UAVs and many other ducted fan based vertical lift systems. The new concept that will significantly reduce the inlet lip separation related performance penalties used a secondary stationary duct system to control "inlet lip separation" occurring especially at elevated forward flight velocities. The (DDF) is self-adjusting in a wide forward flight velocity range. DDFs corrective aerodynamic in influence becomes more pronounced with increasing flight velocity due to its inherent design properties. RANS simulations of the flow around rotor blades and duct geometry in the rotating frame of reference provided a comprehensive description of the tip leakage and passage flow in the flow environment of the two ducted fan research facilities developed throughout this thesis. The aerodynamic measurements and results of the RANS simulation showed good agreement especially near the tip region. A number of novel tip treatments based on custom designed pressure side extensions were introduced. Various tip leakage mitigation schemes were introduced by varying the chordwise location and the width of the extension in the circumferential direction. The current study showed that a proper selection of the pressure side bump location and width were the two critical parameters in influencing the success of the tip leakage mitigation approach. Significant gains in axial mean velocity component were observed when a proper pressure side tip extension was used. It is also observed that an effective tip leakage mitigation scheme significantly reduced the tangential velocity component near the tip of the axial fan blade. Reduced tip clearance related flow interactions were essential in improving the energy efficiency and range of ducted fan based vehicle. Full and inclined pressure side tip squealers were designed. Squealer tips were effective in changing the overall trajectory of the tip vortex to a higher path in radial direction. The interaction of rotor blades and tip vortex was effectively reduced and aerodynamic performance of the rotor blades was improved. The overall aerodynamic gain was a measurable reduction in leakage mass flow rate. This leakage reduction increased thrust significantly. Full and inclined pressure side tip squealers increased thrust obtained in hover condition by 9.1 % and 9.6 % respectively. A reduction or elimination of the momentum deficit in tip vortices is essential to reduce the adverse performance effects originating from the unsteady and highly turbulent tip leakage flows rotating against a stationary casing. The novel tip treatments developed throughout this thesis research are highly effective in reducing the adverse performance effects of ducted fan systems developed for VTOL vehicles. (Abstract shortened by UMI.)

Investigations on the Aerodynamic Characteristics and Blade Excitations of the Radial Turbine with Pulsating Inlet Flow

NASA Astrophysics Data System (ADS)

Liu, Yixiong; Yang, Ce; Yang, Dengfeng; Zhang, Rui

2016-04-01

The aerodynamic performance, detailed unsteady flow and time-based excitations acting on blade surfaces of a radial flow turbine have been investigated with pulsation flow condition. The results show that the turbine instantaneous performance under pulsation flow condition deviates from the quasi-steady value significantly and forms obvious hysteretic loops around the quasi-steady conditions. The detailed analysis of unsteady flow shows that the characteristic of pulsation flow field in radial turbine is highly influenced by the pulsation inlet condition. The blade torque, power and loading fluctuate with the inlet pulsation wave in a pulse period. For the blade excitations, the maximum and the minimum blade excitations conform to the wave crest and wave trough of the inlet pulsation, respectively, in time-based scale. And toward blade chord direction, the maximum loading distributes along the blade leading edge until 20% chord position and decreases from the leading to trailing edge.

Verification Assessment of Flow Boundary Conditions for CFD Analysis of Supersonic Inlet Flows

NASA Technical Reports Server (NTRS)

Slater, John W.

2002-01-01

Boundary conditions for subsonic inflow, bleed, and subsonic outflow as implemented into the WIND CFD code are assessed with respect to verification for steady and unsteady flows associated with supersonic inlets. Verification procedures include grid convergence studies and comparisons to analytical data. The objective is to examine errors, limitations, capabilities, and behavior of the boundary conditions. Computational studies were performed on configurations derived from a "parameterized" supersonic inlet. These include steady supersonic flows with normal and oblique shocks, steady subsonic flow in a diffuser, and unsteady flow with the propagation and reflection of an acoustic disturbance.

Harmonic uniflow engine

DOEpatents

Bennett, Charles L.

2016-03-22

A reciprocating-piston uniflow engine includes a harmonic oscillator inlet valve capable of oscillating at a resonant frequency for controlling the flow of working fluid into the engine. In particular, the inlet valve includes an inlet valve head and a spring arranged together as a harmonic oscillator so that the inlet valve head is moveable from an unbiased equilibrium position to a biased closed position occluding an inlet. When released, the inlet valve head undergoes a single oscillation past the equilibrium position to a maximum open position and returns to a biased return position close to the closed position to choke the flow and produce a pressure drop across the inlet valve causing the inlet valve to close. In other embodiments, the harmonic oscillator arrangement of the inlet valve enables the uniflow engine to be reversibly operated as a uniflow compressor.

Turboprop engine and method of operating the same

DOE Office of Scientific and Technical Information (OSTI.GOV)

Klees, G.W.; Johnson, P.E.

1986-02-11

This patent describes a turboprop engine consisting of: 1.) A compressor; 2.) A turbine; 3.) A combustion section; 4.) A variable pitch propeller; 5.) A speed reducing transmission; 6.) An air inlet; 7.) An air inlet bypass; 8.) An air outlet bypass duct; 9.) A flow control operatively positioned to receive air flow from the air inlet bypass and air flow from the low pressure compressor component. To direct the air flow to the air outlet bypass duct, and the air flow to the high pressure compressor component, the flow control has a first position where the air flow ismore » from. The high and low pressure compressor components and is directed to the air outlet bypass duct. The flow control has a second position for the air flow from the air inlet bypass duct to the air outlet bypass duct and air from the low pressure compressor component is directed to the high pressure compressor component. A method of operating a turboprop engine.« less

Estimation of additive forces and moments for supersonic inlets

NASA Technical Reports Server (NTRS)

Perkins, Stanley C., Jr.; Dillenius, Marnix F. E.

1991-01-01

A technique for estimating the additive forces and moments associated with supersonic, external compression inlets as a function of mass flow ratio has been developed. The technique makes use of a low order supersonic paneling method for calculating minimum additive forces at maximum mass flow conditions. A linear relationship between the minimum additive forces and the maximum values for fully blocked flow is employed to obtain the additive forces at a specified mass flow ratio. The method is applicable to two-dimensional inlets at zero or nonzero angle of attack, and to axisymmetric inlets at zero angle of attack. Comparisons with limited available additive drag data indicate fair to good agreement.

Development of a three-dimensional supersonic inlet flow analysis

NASA Technical Reports Server (NTRS)

Buggeln, R. C.; Mcdonald, H.; Levy, R.; Kreskovsky, J. P.

1980-01-01

A method for computing three dimensional flow in supersonic inlets is described. An approximate set of governing equations is given for viscous flows which have a primary flow direction. The governing equations are written in general orthogonal coordinates. These equations are modified in the subsonic region of the flow to prevent the phenomenon of branching. Results are presented for the two sample cases: a Mach number equals 2.5 flow in a square duct, and a Mach number equals 3.0 flow in a research jet engine inlet. In the latter case the computed results are compared with the experimental data. A users' manual is included.

Modular hybrid plasma reactor and related systems and methods

DOEpatents

Kong, Peter C.; Grandy, Jon D.; Detering, Brent A.

2010-06-22

A device, method and system for generating a plasma is disclosed wherein an electrical arc is established and the movement of the electrical arc is selectively controlled. In one example, modular units are coupled to one another to collectively define a chamber. Each modular unit may include an electrode and a cathode spaced apart and configured to generate an arc therebetween. A device, such as a magnetic or electromagnetic device, may be used to selectively control the movement of the arc about a longitudinal axis of the chamber. The arcs of individual modules may be individually controlled so as to exhibit similar or dissimilar motions about the longitudinal axis of the chamber. In another embodiment, an inlet structure may be used to selectively define the flow path of matter introduced into the chamber such that it travels in a substantially circular or helical path within the chamber.

Light absorption cell combining variable path and length pump

DOEpatents

Prather, W.S.

1993-12-07

A device is described for use in making spectrophotometric measurements of fluid samples. In particular, the device is a measurement cell containing a movable and a fixed lens with a sample of the fluid there between and through which light shines. The cell is connected to a source of light and a spectrophotometer via optic fibers. Movement of the lens varies the path length and also pumps the fluid into and out of the cell. Unidirectional inlet and exit valves cooperate with the movable lens to assure a one-way flow of fluid through the cell. A linear stepper motor controls the movement of the lens and cycles it from a first position closer to the fixed lens and a second position farther from the fixed lens, preferably at least 10 times per minute for a nearly continuous stream of absorption spectrum data. 2 figures.

Comparisons of a Three-Dimensional, Full Navier Stokes Computer Model with High Mach Number Combuster Test Data

NASA Technical Reports Server (NTRS)

Watkins, William B.

1990-01-01

Comparisons between scramjet combustor data and a three-dimensional full Navier-Stokes calculation have been made to verify and substantiate computational fluid dynamics (CFD) codes and application procedures. High Mach number scramjet combustor development will rely heavily on CFD applications to provide wind tunnel-equivalent data of quality sufficient to design, build and fly hypersonic aircraft. Therefore. detailed comparisons between CFD results and test data are imperative. An experimental case is presented, for which combustor wall static pressures were measured and flow-fieid interferograms were obtained. A computer model was done of the experiment, and counterpart parameters are compared with experiment. The experiment involved a subscale combustor designed and fabricated for the National Aero-Space Plane Program, and tested in the Calspan Corporation 96" hypersonic shock tunnel. The combustor inlet ramp was inclined at a 20 angle to the shock tunnel nozzle axis, and resulting combustor entrance flow conditions simulated freestream M=10. The combustor body and cowl walls were instrumented with static pressure transducers, and the combustor lateral walls contained windows through which flowfield holographic interferograms were obtained. The CFD calculation involved a three-dimensional time-averaged full Navier-Stokes code applied to the axial flow segment containing fuel injection and combustion. The full Navier-Stokes approach allowed for mixed supersonic and subsonic flow, downstream-upstream communication in subsonic flow regions, and effects of adverse pressure gradients. The code included hydrogen-air chemistry in the combustor segment which begins near fuel injection and continues through combustor exhaust. Combustor ramp and inlet segments on the combustor lateral centerline were modelled as two dimensional. Comparisons to be shown include calculated versus measured wall static pressures as functions of axial flow coordinate, and calculated path-averaged density contours versus an holographic Interferogram.

Active unsteady aerodynamic suppression of rotating stall in an incompressible flow centrifugal compressor with vaned diffuser

NASA Technical Reports Server (NTRS)

Lawless, Patrick B.; Fleeter, Sanford

1991-01-01

A mathematical model is developed to analyze the suppression of rotating stall in an incompressible flow centrifugal compressor with a vaned diffuser, thereby addressing the important need for centrifugal compressor rotating stall and surge control. In this model, the precursor to to instability is a weak rotating potential velocity perturbation in the inlet flow field that eventually develops into a finite disturbance. To suppress the growth of this potential disturbance, a rotating control vortical velocity disturbance is introduced into the impeller inlet flow. The effectiveness of this control is analyzed by matching the perturbation pressure in the compressor inlet and exit flow fields with a model for the unsteady behavior of the compressor. To demonstrate instability control, this model is then used to predict the control effectiveness for centrifugal compressor geometries based on a low speed research centrifugal compressor. These results indicate that reductions of 10 to 15 percent in the mean inlet flow coefficient at instability are possible with control waveforms of half the magnitude of the total disturbance at the inlet.

Large-Scale Low-Boom Inlet Test Overview

NASA Technical Reports Server (NTRS)

Hirt, Stefanie

2011-01-01

This presentation provides a high level overview of the Large-Scale Low-Boom Inlet Test and was presented at the Fundamental Aeronautics 2011 Technical Conference. In October 2010 a low-boom supersonic inlet concept with flow control was tested in the 8'x6' supersonic wind tunnel at NASA Glenn Research Center (GRC). The primary objectives of the test were to evaluate the inlet stability and operability of a large-scale low-boom supersonic inlet concept by acquiring performance and flowfield validation data, as well as evaluate simple, passive, bleedless inlet boundary layer control options. During this effort two models were tested: a dual stream inlet intended to model potential flight hardware and a single stream design to study a zero-degree external cowl angle and to permit surface flow visualization of the vortex generator flow control on the internal centerbody surface. The tests were conducted by a team of researchers from NASA GRC, Gulfstream Aerospace Corporation, University of Illinois at Urbana-Champaign, and the University of Virginia

Some flow phenomena in a constant area duct with a Borda type inlet including the critical region

NASA Technical Reports Server (NTRS)

Hendricks, R. C.; Simoneau, R. J.

1978-01-01

Mass limiting flow characteristics for a 55 L/D tube with a Borda type inlet were assessed over large ranges of temperature and pressure, using fluid nitrogen. Under certain conditions, separation and pressure drop at the inlet was sufficiently strong to permit partial vaporization and the remaining fluid flowed through the tube as if it were a free jet. An empirical relation was determined which defines conditions under which this type of flow can occur. A flow coefficient is presented which enables estimations of flow rates over the experimental range. A flow rate stagnation pressure map for selected stagnation isotherms and pressure profiles document these flow phenomena.

Apparatus and method for continuous production of materials

DOEpatents

Chang, Chih-hung; Jin, Hyungdae

2014-08-12

Embodiments of a continuous-flow injection reactor and a method for continuous material synthesis are disclosed. The reactor includes a mixing zone unit and a residence time unit removably coupled to the mixing zone unit. The mixing zone unit includes at least one top inlet, a side inlet, and a bottom outlet. An injection tube, or plurality of injection tubes, is inserted through the top inlet and extends past the side inlet while terminating above the bottom outlet. A first reactant solution flows in through the side inlet, and a second reactant solution flows in through the injection tube(s). With reference to nanoparticle synthesis, the reactant solutions combine in a mixing zone and form nucleated nanoparticles. The nucleated nanoparticles flow through the residence time unit. The residence time unit may be a single conduit, or it may include an outer housing and a plurality of inner tubes within the outer housing.

A study of the compressible flow through a diffusing S-duct

NASA Technical Reports Server (NTRS)

Wellborn, Steven R.; Okiishi, Theodore H.; Reichert, Bruce A.

1993-01-01

Benchmark aerodynamic data are presented for compressible flow through a representative S-duct configuration. A numerical prediction of the S-duct flow field, obtained from a subsonic parabolized Navier-Stokes algorithm, is also shown. The experimental and numerical results are compared. Measurements of the three-dimensional velocity field, total pressures, and static pressures were obtained at five cross-sectional planes. Aerodynamic data were gathered with calibrated pneumatic probes. Surface static pressure and surface flow visualization data were also acquired. All reported tests were conducted with an inlet centerline Mach number of 0.6. The Reynolds number, based on the inlet centerline velocity and duct inlet diameter, was 2.6 x 10(exp 6). Thin inlet turbulent boundary layers existed. The collected data should be beneficial to aircraft inlet designers and the measurements are suitable for the validation of computational codes. The results show that a region of streamwise flow separation occurred within the duct. Details about the separated flow region, including mechanisms which drive this complicated flow phenomenon, are discussed. Results also indicate that the duct curvature induces strong pressure driven secondary flows. The cross flows evolve into counter-rotating vortices. These vortices convect low momentum fluid of the boundary layer toward the center of the duct, degrading both the uniformity and magnitude of the total pressure profile.

Harmonic engine

DOEpatents

Bennett, Charles L.; Sewall, Noel; Boroa, Carl

2014-08-19

An engine based on a reciprocating piston engine that extracts work from pressurized working fluid. The engine includes a harmonic oscillator inlet valve capable of oscillating at a resonant frequency for controlling the flow of working fluid into of the engine. In particular, the inlet valve includes an inlet valve head and a spring arranged together as a harmonic oscillator so that the inlet valve head is moveable from an unbiased equilibrium position to a biased closed position occluding an inlet. Upon releasing the inlet valve the inlet valve head undergoes a single oscillation past the equilibrium positio to a maximum open position and returns to a biased return position close to the closed position to choke the flow and produce a pressure drop across the inlet valve causing the inlet valve to close. Protrusions carried either by the inlet valve head or piston head are used to bump open the inlet valve from the closed position and initiate the single oscillation of the inlet valve head, and protrusions carried either by the outlet valve head or piston head are used to close the outlet valve ahead of the bump opening of the inlet valve.

Force and pressure-recovery characteristics at supersonic speeds of a conical nose inlet with bypasses discharging outward from the body axis

NASA Technical Reports Server (NTRS)

Beke, Andrew; Allen, J L

1953-01-01

Aerodynamic and performance characteristics of a conical spike nacelle-type inlet with two bypasses are presented at Mach numbers of 1.6, 1.8, and 2.0 for angles of attach up to 90 degrees. The bypasses were located 6 inlet diameters downstream of the inlet and were designed to discharge the bypass mass flow outward from the body axis. The inlet was designed to attain a mass-flow ratio of unity at a Mach number of 2.0. It is shown that discharging the bypass mass flow outward from the body nearly doubles the critical drag of a similar configuration but with bypass discharge in an axial direction. As a result of this greater drag, the net force on the model in the flight direction is reduced when comparison is made with the axial discharge case. The lift and pitching-moment coefficients are slightly higher than those for a configuration without bypasses. Approximately 25 % of the maximum inlet mass flow was discharged through the bypasses, and the pressure-recovery and mass-flow characteristics were in qualitative and quantitative agreement with the results of an investigation of a similar configuration with axial discharge.

3-D viscous flow CFD analysis of the propeller effect on an advanced ducted propeller subsonic inlet

NASA Technical Reports Server (NTRS)

Iek, Chanthy; Boldman, Donald R.; Ibrahim, Mounir

1993-01-01

The time-marching Navier-Stokes code PARC3D was used to study the 3D viscous flow associated with an advanced ducted propeller subsonic inlet at take-off operating conditions. At a free stream Mach number of 0.2, experimental data for the inlet-with-propeller test model indicated that the airflow was attached on the cowl windward lip at an angle of attack of 25 deg became unstable at 29 deg, and separated at 30 deg. An experimental study with a similar inlet and without propeller (through-flow) indicated that flow separation occurred at an angle of attack a few degrees below the value observed when the inlet was tested with the propeller, indicating the propeller's favorable effect on inlet performance. In the present numerical study, flow blockage analogous to the propeller was modeled via a PARC3D computational boundary condition (BC), the 'screen BC', based on 1-1/2 dimension actuator disk theory. The application of the screen BC in this numerical study provided results similar to those of past experimental efforts in which either the blockage device or the propeller was used.

Osborne Reynolds pipe flow: direct numerical simulation from laminar to fully-developed turbulence

NASA Astrophysics Data System (ADS)

Adrian, R. J.; Wu, X.; Moin, P.; Baltzer, J. R.

2014-11-01

Osborne Reynolds' pipe experiment marked the onset of modern viscous flow research, yet the detailed mechanism carrying the laminar state to fully-developed turbulence has been quite elusive, despite notable progress related to dynamic edge-state theory. Here, we continue our direct numerical simulation study on this problem using a 250R long, spatially-developing pipe configuration with various Reynolds numbers, inflow disturbances, and inlet base flow states. For the inlet base flow, both fully-developed laminar profile and the uniform plug profile are considered. Inlet disturbances consist of rings of turbulence of different width and radial location. In all the six cases examined so far, energy norms show exponential growth with axial distance until transition after an initial decay near the inlet. Skin-friction overshoots the Moody's correlation in most, but not all, the cases. Another common theme is that lambda vortices amplified out of susceptible elements in the inlet disturbances trigger rapidly growing hairpin packets at random locations and times, after which infant turbulent spots appear. Mature turbulent spots in the pipe transition are actually tight concentrations of hairpin packets looking like a hairpin forest. The plug flow inlet profile requires much stronger disturbances to transition than the parabolic profile.

Tapered plug foam spray apparatus

NASA Technical Reports Server (NTRS)

Allen, Peter B. (Inventor)

1996-01-01

A two-component foam spray gun is readily disassembled for cleaning. It includes a body (1) with reactant (12, 14) and purge gas (16) inlet ports. A moldable valve packing (32) inside the body has a tapered conical interior surface (142), and apertures which match the reactant ports. A valve/tip (40) has a conical outer surface (48) which mates with the valve packing (32). The valve/tip (40) is held in place by a moldable packing washer (34), held at non-constant pressure by a screw (36, 38). The interior of the valve/tip (40) houses a removable mixing chamber (50). The mixing chamber (50) has direct flow orifices (60) and an auxiliary flow path (58, 60) which ameliorate pressure surges. The spray gun can be disassembled for cleaning without disturbing the seal, by removing the valve/tip (40) to the rear, thereby breaking it free of the conical packing. Rotation of the valve/tip (40) relative to the body (1) shuts off the reactant flow, and starts the purge gas flow.

Cavitation in centrifugal pump with rotating walls of axial inlet device

NASA Astrophysics Data System (ADS)

Moloshnyi, O.; Sotnyk, M.

2017-08-01

The article deals with the analysis of cavitation processes in the flowing part of the double entry centrifugal pump. The analysis is conducted using numerical modeling of the centrifugal pump operating process in the software environment ANSYS CFX. Two models of the axial inlet device is researched. It is shaped by a cylindrical section and diffuser section in front of the impeller, which includes fairing. The walls of the axial inlet device rotate with the same speed as the pump rotor. The numerical experiment is conducted under the condition of the flow rate change and absolute pressure at the inlet. The analysis shows that the pump has the average statistical cavitation performance. The occurrence of the cavitation in the axial inlet device is after narrowing the cross-section of flow channel and at the beginning of the diffuser section. Additional sudden expansion at the outlet of the axial inlet diffuser section does not affect the cavitation characteristics of the impeller, however, improves cavitation characteristics of the axial inlet device. For considered geometric parameters of the axial inlet device the cavitation in the impeller begins earlier than in the axial inlet device. That is, the considered design of the axial inlet device will not be subjected to destruction at the ensuring operation without cavitation in the impeller.

Wind-tunnel tests on a 3-dimensional fixed-geometry scramjet inlet at M = 2.30 to 4.60

NASA Technical Reports Server (NTRS)

Mueller, J. N.; Trexler, C. A.; Souders, S. W.

1977-01-01

Wind-tunnel tests were conducted on a baseline scramjet inlet model having fixed geometry and swept leading edges at M = 2.30, 2.96, 3.95, and 4.60 in the Langley unitary plan wind tunnel. The unit Reynolds number of the tests was held constant at 6.56 million per meter (2 million per foot). The objectives of the tests were to establish inlet performance and starting characteristics in the lower Mach number range of operation (less than M = 5). Surface pressures obtained on the inlet components are presented, along with the results of the internal flow surveys made at the throat and capture stations of the inlet. Contour plots of the inlet-flow-field parameters such as Mach numbers, pressure recovery, flow capture, local static and total pressure ratios at the survey stations are shown for the test Mach numbers.

Opposed-flow virtual cyclone for particle concentration

DOEpatents

Rader, Daniel J.; Torczynski, John R.

2000-12-05

An opposed-flow virtual cyclone for aerosol collation which can accurately collect, classify, and concentrate (enrich) particles in a specific size range. The opposed-flow virtual cyclone is a variation on the virtual cyclone and has its inherent advantages (no-impact particle separation in a simple geometry), while providing a more robust design for concentrating particles in a flow-through type system. The opposed-flow virtual cyclone consists of two geometrically similar virtual cyclones arranged such that their inlet jets are inwardly directed and symmetrically opposed relative to a plane of symmetry located between the two inlet slits. A top plate bounds both jets on the "top" side of the inlets, while the other or lower wall curves "down" and away from each inlet jet. Each inlet jet will follow the adjacent lower wall as it turns away, and that particles will be transferred away from the wall and towards the symmetry plane by centrifugal action. After turning, the two jets merge smoothly along the symmetry line and flow parallel to it through the throat. Particles are transferred from the main flows, across a dividing streamline, and into a central recirculating region, where particle concentrations become greatly increased relative to the main stream.

On the question of starting conditions for frontal axisymmetric inlets tested in hot-shot wind tunnels

NASA Astrophysics Data System (ADS)

Gounko, Yu. P.; Mazhul, I. I.

2017-05-01

The work presents the results of an analysis of starting conditions for some frontal axisymmetric inlets of internal compression tested at freestream Mach numbers M = 3-8.4 in the hot-shot wind tunnels based at Khristianovich Institute of Theoretical and Applied Mechanics (ITAM). The results of these inlets test are compared with the data of numerical computations of inviscid, laminar, and turbulent flows carried out by the pseudo-unsteady method. There were determined the inlet throat areas limiting either with regard to the inlet starting or with regard to providing the maximally possible degree of geometric compression of the inlet-captured supersonic airstream at its deceleration in the already started inlet. Reshaping of computed flow patterns in the inlets depending on the variation of the minimal cross section of the inlet internal duct is analyzed.

Overview of the Turbine Based Combined Cycle Discipline

NASA Technical Reports Server (NTRS)

Thomas, Scott R.; Walker, James F.; Pittman, James L.

2009-01-01

The NASA Fundamental Aeronautics Hypersonics project is focused on technologies for combined cycle, airbreathing propulsions systems to enable reusable launch systems for access to space. Turbine Based Combined Cycle (TBCC) propulsion systems offer specific impulse (Isp) improvements over rocket-based propulsion systems in the subsonic takeoff and return mission segments and offer improved safety. The potential to realize more aircraft-like operations with expanded launch site capability and reduced system maintenance are additional benefits. The most critical TBCC enabling technologies as identified in the National Aeronautics Institute (NAI) study were: 1) mode transition from the low speed propulsion system to the high speed propulsion system, 2) high Mach turbine engine development, 3) transonic aero-propulsion performance, 4) low-Mach-number dual-mode scramjet operation, 5) innovative 3-D flowpath concepts and 6) innovative turbine based combined cycle integration. To address several of these key TBCC challenges, NASA s Hypersonics project (TBCC Discipline) initiated an experimental mode transition task that includes an analytic research endeavor to assess the state-of-the-art of propulsion system performance and design codes. This initiative includes inlet fluid and turbine performance codes and engineering-level algorithms. This effort has been focused on the Combined Cycle Engine Large-Scale Inlet Mode Transition Experiment (CCE LIMX) which is a fully integrated TBCC propulsion system with flow path sizing consistent with previous NASA and DoD proposed Hypersonic experimental flight test plans. This experiment is being tested in the NASA-GRC 10 x 10 Supersonic Wind Tunnel (SWT) Facility. The goal of this activity is to address key hypersonic combined-cycle-engine issues: (1) dual integrated inlet operability and performance issues unstart constraints, distortion constraints, bleed requirements, controls, and operability margins, (2) mode-transition constraints imposed by the turbine and the ramjet/scramjet flow paths (imposed variable geometry requirements), (3) turbine engine transients (and associated time scales) during transition, (4) high-altitude turbine engine re-light, and (5) the operating constraints of a Mach 3-7 combustor (specific to the TBCC). The model will be tested in several test phases to develop a unique TBCC database to assess and validate design and analysis tools and address operability, integration, and interaction issues for this class of advanced propulsion systems. The test article and all support equipment is complete and available at the facility. The test article installation and facility build-up in preparation for the inlet performance and operability characterization is near completion and testing is planned to commence in FY11.

Computational study of the effects of shroud geometric variation on turbine performance in a 1.5-stage high-loaded turbine

NASA Astrophysics Data System (ADS)

Jia, Wei; Liu, Huoxing

2013-10-01

Generally speaking, main flow path of gas turbine is assumed to be perfect for standard 3D computation. But in real engine, the turbine annulus geometry is not completely smooth for the presence of the shroud and associated cavity near the end wall. Besides, shroud leakage flow is one of the dominant sources of secondary flow in turbomachinery, which not only causes a deterioration of useful work but also a penalty on turbine efficiency. It has been found that neglect shroud leakage flow makes the computed velocity profiles and loss distribution significantly different to those measured. Even so, the influence of shroud leakage flow is seldom taken into consideration during the routine of turbine design due to insufficient understanding of its impact on end wall flows and turbine performance. In order to evaluate the impact of tip shroud geometry on turbine performance, a 3D computational investigation for 1.5-stage turbine with shrouded blades was performed in this paper. The following geometry parameters were varied respectively: Inlet cavity length and exit cavity length

Numerical Modeling of Active Flow Control in a Boundary Layer Ingesting Offset Inlet

NASA Technical Reports Server (NTRS)

Allan, Brian G.; Owens, Lewis R.; Berrier, Bobby L.

2004-01-01

This investigation evaluates the numerical prediction of flow distortion and pressure recovery for a boundary layer ingesting offset inlet with active flow control devices. The numerical simulations are computed using a Reynolds averaged Navier-Stokes code developed at NASA. The numerical results are validated by comparison to experimental wind tunnel tests conducted at NASA Langley Research Center at both low and high Mach numbers. Baseline comparisons showed good agreement between numerical and experimental results. Numerical simulations for the inlet with passive and active flow control also showed good agreement at low Mach numbers where experimental data has already been acquired. Numerical simulations of the inlet at high Mach numbers with flow control jets showed an improvement of the flow distortion. Studies on the location of the jet actuators, for the high Mach number case, were conducted to provide guidance for the design of a future experimental wind tunnel test.

Parametric Data from a Wind Tunnel Test on a Rocket-Based Combined-Cycle Engine Inlet

NASA Technical Reports Server (NTRS)

Fernandez, Rene; Trefny, Charles J.; Thomas, Scott R.; Bulman, Mel J.

2001-01-01

A 40-percent scale model of the inlet to a rocket-based combined-cycle (RBCC) engine was tested in the NASA Glenn Research Center 1- by 1-Foot Supersonic Wind Tunnel (SWT). The full-scale RBCC engine is scheduled for test in the Hypersonic Tunnel Facility (HTF) at NASA Glenn's Plum Brook Station at Mach 5 and 6. This engine will incorporate the configuration of this inlet model which achieved the best performance during the present experiment. The inlet test was conducted at Mach numbers of 4.0, 5.0, 5.5, and 6.0. The fixed-geometry inlet consists of an 8 deg.. forebody compression plate, boundary layer diverter, and two compressive struts located within 2 parallel sidewalls. These struts extend through the inlet, dividing the flowpath into three channels. Test parameters investigated included strut geometry, boundary layer ingestion, and Reynolds number (Re). Inlet axial pressure distributions and cross-sectional Pitot-pressure surveys at the base of the struts were measured at varying back-pressures. Inlet performance and starting data are presented. The inlet chosen for the RBCC engine self-started at all Mach numbers from 4 to 6. Pitot-pressure contours showed large flow nonuniformity on the body-side of the inlet. The inlet provided adequate pressure recovery and flow quality for the RBCC cycle even with the flow separation.

Verification of a three-dimensional resin transfer molding process simulation model

NASA Technical Reports Server (NTRS)

Fingerson, John C.; Loos, Alfred C.; Dexter, H. Benson

1995-01-01

Experimental evidence was obtained to complete the verification of the parameters needed for input to a three-dimensional finite element model simulating the resin flow and cure through an orthotropic fabric preform. The material characterizations completed include resin kinetics and viscosity models, as well as preform permeability and compaction models. The steady-state and advancing front permeability measurement methods are compared. The results indicate that both methods yield similar permeabilities for a plain weave, bi-axial fiberglass fabric. Also, a method to determine principal directions and permeabilities is discussed and results are shown for a multi-axial warp knit preform. The flow of resin through a blade-stiffened preform was modeled and experiments were completed to verify the results. The predicted inlet pressure was approximately 65% of the measured value. A parametric study was performed to explain differences in measured and predicted flow front advancement and inlet pressures. Furthermore, PR-500 epoxy resin/IM7 8HS carbon fabric flat panels were fabricated by the Resin Transfer Molding process. Tests were completed utilizing both perimeter injection and center-port injection as resin inlet boundary conditions. The mold was instrumented with FDEMS sensors, pressure transducers, and thermocouples to monitor the process conditions. Results include a comparison of predicted and measured inlet pressures and flow front position. For the perimeter injection case, the measured inlet pressure and flow front results compared well to the predicted results. The results of the center-port injection case showed that the predicted inlet pressure was approximately 50% of the measured inlet pressure. Also, measured flow front position data did not agree well with the predicted results. Possible reasons for error include fiber deformation at the resin inlet and a lag in FDEMS sensor wet-out due to low mold pressures.

Power-on performance predictions for a complete generic hypersonic vehicle configuration

NASA Technical Reports Server (NTRS)

Bennett, Bradford C.

1991-01-01

The Compressible Navier-Stokes (CNS) code was developed to compute external hypersonic flow fields. It has been applied to various hypersonic external flow applications. Here, the CNS code was modified to compute hypersonic internal flow fields. Calculations were performed on a Mach 18 sidewall compression inlet and on the Lewis Mach 5 inlet. The use of the ARC3D diagonal algorithm was evaluated for internal flows on the Mach 5 inlet flow. The initial modifications to the CNS code involved generalization of the boundary conditions and the addition of viscous terms in the second crossflow direction and modifications to the Baldwin-Lomax turbulence model for corner flows.

Inlet flow test calibration for a small axial compressor rig. Part 2: CFD compared with experimental results

NASA Technical Reports Server (NTRS)

Miller, D. P.; Prahst, P. S.

1995-01-01

An axial compressor test rig has been designed for the operation of small turbomachines. A flow test was run to calibrate and determine the source and magnitudes of the loss mechanisms in the compressor inlet for a highly loaded two-stage axial compressor test. Several flow conditions and inlet guide vane (IGV) angle settings were established, for which detailed surveys were completed. Boundary layer bleed was also provided along the casing of the inlet behind the support struts and ahead of the IGV. Several computational fluid dynamics (CFD) calculations were made for selected flow conditions established during the test. Good agreement between the CFD and test data were obtained for these test conditions.

Effect of a dual inlet channel on cell loading in microfluidics.

PubMed

Yun, Hoyoung; Kim, Kisoo; Lee, Won Gu

2014-11-01

Unwanted sedimentation and attachment of a number of cells onto the bottom channel often occur on relatively large-scale inlets of conventional microfluidic channels as a result of gravity and fluid shear. Phenomena such as sedimentation have become recognized problems that can be overcome by performing microfluidic experiments properly, such as by calculating a meaningful output efficiency with respect to real input. Here, we present a dual-inlet design method for reducing cell loss at the inlet of channels by adding a new " upstream inlet " to a single main inlet design. The simple addition of an upstream inlet can create a vertically layered sheath flow prior to the main inlet for cell loading. The bottom layer flow plays a critical role in preventing the cells from attaching to the bottom of the channel entrance, resulting in a low possibility of cell sedimentation at the main channel entrance. To provide proof-of-concept validation, we applied our design to a microfabricated flow cytometer system (μFCS) and compared the cell counting efficiency of the proposed μFCS with that of the previous single-inlet μFCS and conventional FCS. We used human white blood cells and fluorescent microspheres to quantitatively evaluate the rate of cell sedimentation in the main inlet and to measure fluorescence sensitivity at the detection zone of the flow cytometer microchip. Generating a sheath flow as the bottom layer was meaningfully used to reduce the depth of field as well as the relative deviation of targets in the z-direction (compared to the x-y flow plane), leading to an increased counting sensitivity of fluorescent detection signals. Counting results using fluorescent microspheres showed both a 40% reduction in the rate of sedimentation and a 2-fold higher sensitivity in comparison with the single-inlet μFCS. The results of CD4(+) T-cell counting also showed that the proposed design results in a 25% decrease in the rate of cell sedimentation and a 28% increase in sensitivity when compared to the single-inlet μFCS. This method is simple and easy to use in design, yet requires no additional time or cost in fabrication. Furthermore, we expect that this approach could potentially be helpful for calculating exact cell loading and counting efficiency for a small input number of cells, such as primary cells and rare cells, in microfluidic channel applications.

The Total-Pressure Recovery and Drag Characteristics of Several Auxiliary Inlets at Transonic Speeds

NASA Technical Reports Server (NTRS)

Dennard, John S.

1959-01-01

Several flush and scoop-type auxiliary inlets have been tested for a range of Mach numbers from 0.55 to 1.3 to determine their transonic total-pressure recovery and drag characteristics. The inlet dimensions were comparable with the thickness of the boundary layer in which they were tested. Results indicate that flush inlets should be inclined at very shallow angles with respect to the surface for optimum total-pressure recovery and drag characteristics. Deep, narrow inlets have lower drag than wide shallow ones at Mach numbers greater than 0.9 but at lower Mach numbers the wider inlets proved superior. Inlets with a shallow approach ramp, 7 deg, and diverging ramp walls which incorporated boundary-layer bypass had lower drag than any other inlet tested for Mach numbers up to 1.2 and had the highest pressure recovery of all of the flush inlets. The scoop inlets, which operated in a higher velocity flow than the flush inlets, had higher drag coefficients. Several of these auxiliary inlets projected multiple, periodic shock waves into the stream when they were operated at low mass-flow ratios.

The Effect of Upstream Vane Wakes on Annular Diffuser Flows

NASA Astrophysics Data System (ADS)

Cherry, Erica; Padilla, Angelina; Elkins, Christopher; Eaton, John

2008-11-01

Experiments were performed to determine the sensitivity to inlet conditions of the flow in two annular diffusers. One of the diffusers was a conservative design typical of a diffuser directly upstream of the combustor in a jet engine. The other had the same length and inlet shape as the first diffuser but a larger area ratio and was meant to operate on the verge of separation. Each diffuser was connected to two different inlets, one containing a fully-developed channel flow, the other containing wakes from a row of airfoils. Three-component velocity measurements were taken on the flow in each inlet/diffuser combination using Magnetic Resonance Velocimetry. Results will be presented on the 3D velocity fields in the two diffusers and the effect of the airfoil wakes on separation and secondary flows.

Experimental results for a two-dimensional supersonic inlet used as a thrust deflecting nozzle

NASA Technical Reports Server (NTRS)

Johns, Albert L.; Burstadt, Paul L.

1984-01-01

Nearly all supersonic V/STOL aircraft concepts are dependent on the thrust deflecting capability of a nozzle. In one unique concept, referred to as the reverse flow dual fan, not only is there a thrust deflecting nozzle for the fan and core engine exit flow, but because of the way the propulsion system operates during vertical takeoff and landing, the supersonic inlet is also used as a thrust deflecting nozzle. This paper presents results of an experimental study to evaluate the performance of a supersonic inlet used as a thrust deflecting nozzle for this reverse flow dual fan concept. Results are presented in terms of nozzle thrust coefficient and thrust vector angle for a number of inlet/nozzle configurations. Flow visualization and nozzle exit flow survey results are also shown.

Fuel cell membrane hydration and fluid metering

DOEpatents

Jones, Daniel O.; Walsh, Michael M.

1999-01-01

A hydration system includes fuel cell fluid flow plate(s) and injection port(s). Each plate has flow channel(s) with respective inlet(s) for receiving respective portion(s) of a given stream of reactant fluid for a fuel cell. Each injection port injects a portion of liquid water directly into its respective flow channel in order to mix its respective portion of liquid water with the corresponding portion of the stream. This serves to hydrate at least corresponding part(s) of a given membrane of the corresponding fuel cell(s). The hydration system may be augmented by a metering system including flow regulator(s). Each flow regulator meters an injecting at inlet(s) of each plate of respective portions of liquid into respective portion(s) of a given stream of fluid by corresponding injection port(s).

Radial lean direct injection burner

DOEpatents

Khan, Abdul Rafey; Kraemer, Gilbert Otto; Stevenson, Christian Xavier

2012-09-04

A burner for use in a gas turbine engine includes a burner tube having an inlet end and an outlet end; a plurality of air passages extending axially in the burner tube configured to convey air flows from the inlet end to the outlet end; a plurality of fuel passages extending axially along the burner tube and spaced around the plurality of air passage configured to convey fuel from the inlet end to the outlet end; and a radial air swirler provided at the outlet end configured to direct the air flows radially toward the outlet end and impart swirl to the air flows. The radial air swirler includes a plurality of vanes to direct and swirl the air flows and an end plate. The end plate includes a plurality of fuel injection holes to inject the fuel radially into the swirling air flows. A method of mixing air and fuel in a burner of a gas turbine is also provided. The burner includes a burner tube including an inlet end, an outlet end, a plurality of axial air passages, and a plurality of axial fuel passages. The method includes introducing an air flow into the air passages at the inlet end; introducing a fuel into fuel passages; swirling the air flow at the outlet end; and radially injecting the fuel into the swirling air flow.

Inlet Flow Test Calibration for a Small Axial Compressor Facility. Part 1: Design and Experimental Results

NASA Technical Reports Server (NTRS)

Miller, D. P.; Prahst, P. S.

1994-01-01

An axial compressor test rig has been designed for the operation of small turbomachines. The inlet region consisted of a long flowpath region with two series of support struts and a flapped inlet guide vane. A flow test was run to calibrate and determine the source and magnitudes of the loss mechanisms in the inlet for a highly loaded two-stage axial compressor test. Several flow conditions and IGV angle settings were established in which detailed surveys were completed. Boundary layer bleed was also provided along the casing of the inlet behind the support struts and ahead of the IGV. A detailed discussion of the flowpath design along with a summary of the experimental results are provided in Part 1.

Impeller leakage flow modeling for mechanical vibration control

NASA Technical Reports Server (NTRS)

Palazzolo, Alan B.

1996-01-01

HPOTP and HPFTP vibration test results have exhibited transient and steady characteristics which may be due to impeller leakage path (ILP) related forces. For example, an axial shift in the rotor could suddenly change the ILP clearances and lengths yielding dynamic coefficient and subsequent vibration changes. ILP models are more complicated than conventional-single component-annular seal models due to their radial flow component (coriolis and centrifugal acceleration), complex geometry (axial/radial clearance coupling), internal boundary (transition) flow conditions between mechanical components along the ILP and longer length, requiring moment as well as force coefficients. Flow coupling between mechanical components results from mass and energy conservation applied at their interfaces. Typical components along the ILP include an inlet seal, curved shroud, and an exit seal, which may be a stepped labyrinth type. Von Pragenau (MSFC) has modeled labyrinth seals as a series of plain annular seals for leakage and dynamic coefficient prediction. These multi-tooth components increase the total number of 'flow coupled' components in the ILP. Childs developed an analysis for an ILP consisting of a single, constant clearance shroud with an exit seal represented by a lumped flow-loss coefficient. This same geometry was later extended to include compressible flow. The objective of the current work is to: supply ILP leakage-force impedance-dynamic coefficient modeling software to MSFC engineers, base on incompressible/compressible bulk flow theory; design the software to model a generic geometry ILP described by a series of components lying along an arbitrarily directed path; validate the software by comparison to available test data, CFD and bulk models; and develop a hybrid CFD-bulk flow model of an ILP to improve modeling accuracy within practical run time constraints.

Boosting devices with integral features for recirculating exhaust gas

DOEpatents

Wu, Ko -Jen

2015-09-15

According to one embodiment of the invention, a compressor housing includes a compressor inlet in fluid communication with a compressor volute configured to house a compressor wheel, the compressor inlet configured to provide a first air flow to the compressor wheel and a compressor outlet in fluid communication with the compressor volute, the compressor outlet configured to direct a compressed gas to an intake manifold. The compressor housing further includes an exhaust gas recirculation inlet port in fluid communication with the compressor volute, the exhaust gas recirculation inlet port being configured to combine an exhaust gas flow with the air flow to the compressor wheel.

A numerical study of automotive turbocharger mixed flow turbine inlet geometry for off design performance

NASA Astrophysics Data System (ADS)

Leonard, T.; Spence, S.; Early, J.; Filsinger, D.

2013-12-01

Mixed flow turbines represent a potential solution to the increasing requirement for high pressure, low velocity ratio operation in turbocharger applications. While literature exists for the use of these turbines at such operating conditions, there is a lack of detailed design guidance for defining the basic geometry of the turbine, in particular, the cone angle - the angle at which the inlet of the mixed flow turbine is inclined to the axis. This investigates the effect and interaction of such mixed flow turbine design parameters. Computational Fluids Dynamics was initially used to investigate the performance of a modern radial turbine to create a baseline for subsequent mixed flow designs. Existing experimental data was used to validate this model. Using the CFD model, a number of mixed flow turbine designs were investigated. These included studies varying the cone angle and the associated inlet blade angle. The results of this analysis provide insight into the performance of a mixed flow turbine with respect to cone and inlet blade angle.

Growth rate of a penny-shaped crack in hydraulic fracturing of rocks

DOE Office of Scientific and Technical Information (OSTI.GOV)

Abe, H.; Keer, L.M.; Mura, T.

1976-01-01

The deformation and growth of a crack, fractured hydraulically, is investigated when fluid is injected from an inlet into the crack at a constant flow rate. The total flow rate at the inlet is divided as follows: flow rate extracted from an outlet hole; fluid loss rate from the crack surface; and total fluid mass change in the crack. Two cases are considered: (1) inlet flow rate is initially greater than the sum of the outlet flow and fluid loss rates; and (2) the reverse holds true. Ranges are shown for which the crack attains stationary states for given inletmore » flow rate and outlet pressure. For these two cases reasonable outlet flow rates are obtained when the outlet pressure is less than or equal to the difference between the tectonic stress and the fluid head at the inlet. Results are expected to be of use in considerations of heat extraction from hot, dry rock.« less

Inlet Unstart Propulsion Integration Wind Tunnel Test Program Completed for High-Speed Civil Transport

NASA Technical Reports Server (NTRS)

Porro, A. Robert

2000-01-01

One of the propulsion system concepts to be considered for the High-Speed Civil Transport (HSCT) is an underwing, dual-propulsion, pod-per-wing installation. Adverse transient phenomena such as engine compressor stall and inlet unstart could severely degrade the performance of one of these propulsion pods. The subsequent loss of thrust and increased drag could cause aircraft stability and control problems that could lead to a catastrophic accident if countermeasures are not in place to anticipate and control these detrimental transient events. Aircraft system engineers must understand what happens during an engine compressor stall and inlet unstart so that they can design effective control systems to avoid and/or alleviate the effects of a propulsion pod engine compressor stall and inlet unstart. The objective of the Inlet Unstart Propulsion Airframe Integration test program was to assess the underwing flow field of a High-Speed Civil Transport propulsion system during an engine compressor stall and subsequent inlet unstart. Experimental research testing was conducted in the 10- by 10-Foot Supersonic Wind Tunnel at the NASA Glenn Research Center at Lewis Field. The representative propulsion pod consisted of a two-dimensional, bifurcated inlet mated to a live turbojet engine. The propulsion pod was mounted below a large flat plate that acted as a wing simulator. Because of the plate s long length (nominally 10-ft wide by 18-ft long), realistic boundary layers could form at the inlet cowl plane. Transient instrumentation was used to document the aerodynamic flow-field conditions during an unstart sequence. Acquiring these data was a significant technical challenge because a typical unstart sequence disrupts the local flow field for about only 50 msec. Flow surface information was acquired via static pressure taps installed in the wing simulator, and intrusive pressure probes were used to acquire flow-field information. These data were extensively analyzed to determine the impact of the unstart transient on the surrounding flow field. This wind tunnel test program was a success, and for the first time, researchers acquired flow-field aerodynamic data during a supersonic propulsion system engine compressor stall and inlet unstart sequence. In addition to obtaining flow-field pressure data, Glenn researchers determined other properties such as the transient flow angle and Mach number. Data are still being reduced, and a comprehensive final report will be released during calendar year 2000.

Hydrodynamic metamaterials: Microfabricated arrays to steer, refract, and focus streams of biomaterials

PubMed Central

Morton, Keith J.; Loutherback, Kevin; Inglis, David W.; Tsui, Ophelia K.; Sturm, James C.; Chou, Stephen Y.; Austin, Robert H.

2008-01-01

We show that it is possible to direct particles entrained in a fluid along trajectories much like rays of light in classical optics. A microstructured, asymmetric post array forms the core hydrodynamic element and is used as a building block to construct microfluidic metamaterials and to demonstrate refractive, focusing, and dispersive pathways for flowing beads and cells. The core element is based on the concept of deterministic lateral displacement where particles choose different paths through the asymmetric array based on their size: Particles larger than a critical size are displaced laterally at each row by a post and move along the asymmetric axis at an angle to the flow, while smaller particles move along streamline paths. We create compound elements with complex particle handling modes by tiling this core element using multiple transformation operations; we show that particle trajectories can be bent at an interface between two elements and that particles can be focused into hydrodynamic jets by using a single inlet port. Although particles propagate through these elements in a way that strongly resembles light rays propagating through optical elements, there are unique differences in the paths of our particles as compared with photons. The unusual aspects of these modular, microfluidic metamaterials form a rich design toolkit for mixing, separating, and analyzing cells and functional beads on-chip. PMID:18495920

Evaluation of range and distortion tolerance for high Mach number transonic fan stages. Task 2: Performance of a 1500-foot-per-second tip speed transonic fan stage with variable geometry inlet guide vanes and stator

NASA Technical Reports Server (NTRS)

Bilwakesh, K. R.; Koch, C. C.; Prince, D. C.

1972-01-01

A 0.5 hub/tip radius ratio compressor stage consisting of a 1500 ft/sec tip speed rotor, a variable camber inlet guide vane and a variable stagger stator was designed and tested with undistorted inlet flow, flow with tip radial distortion, and flow with 90 degrees, one-per-rev, circumferential distortion. At the design speed and design IGV and stator setting the design stage pressure ratio was achieved at a weight within 1% of the design flow. Analytical results on rotor tip shock structure, deviation angle and part-span shroud losses at different operating conditions are presented. The variable geometry blading enabled efficient operation with adequate stall margin at the design condition and at 70% speed. Closing the inlet guide vanes to 40 degrees changed the speed-versus-weight flow relationship along the stall line and thus provided the flexibility of operation at off-design conditions. Inlet flow distortion caused considerable losses in peak efficiency, efficiency on a constant throttle line through design pressure ratio at design speed, stall pressure ratio, and stall margin at the 0 degrees IGV setting and high rotative speeds. The use of the 40 degrees inlet guide vane setting enabled partial recovery of the stall margin over the standard constant throttle line.

Estimating Engine Airflow in Gas-Turbine Powered Aircraft with Clean and Distorted Inlet Flows

NASA Technical Reports Server (NTRS)

Williams, J. G.; Steenken, W. G.; Yuhas, A. J.

1996-01-01

The P404-GF-400 Powered F/A-18A High Alpha Research Vehicle (HARV) was used to examine the impact of inlet-generated total-pressure distortion on estimating levels of engine airflow. Five airflow estimation methods were studied. The Reference Method was a fan corrected airflow to fan corrected speed calibration from an uninstalled engine test. In-flight airflow estimation methods utilized the average, or individual, inlet duct static- to total-pressure ratios, and the average fan-discharge static-pressure to average inlet total-pressure ratio. Correlations were established at low distortion conditions for each method relative to the Reference Method. A range of distorted inlet flow conditions were obtained from -10 deg. to +60 deg. angle of attack and -7 deg. to +11 deg. angle of sideslip. The individual inlet duct pressure ratio correlation resulted in a 2.3 percent airflow spread for all distorted flow levels with a bias error of -0.7 percent. The fan discharge pressure ratio correlation gave results with a 0.6 percent airflow spread with essentially no systematic error. Inlet-generated total-pressure distortion and turbulence had no significant impact on the P404-GE400 engine airflow pumping. Therefore, a speed-flow relationship may provide the best airflow estimate for a specific engine under all flight conditions.

An investigation of the internal and external aerodynamics of cattle trucks

NASA Technical Reports Server (NTRS)

Muirhead, V. U.

1983-01-01

Wind tunnel tests were conducted on a one-tenth scale model of a conventional tractor trailer livestock hauler to determine the air flow through the trailer and the drag of the vehicle. These tests were conducted with the trailer empty and with a full load of simulated cattle. Additionally, the drag was determined for six configurations, of which details for three are documented herein. These are: (1) conventional livestock trailer empty, (2) conventional trailer with smooth sides (i.e., without ventilation openings), and (3) a stream line tractor with modified livestock trailer (cab streamlining and gap fairing). The internal flow of the streamlined modification with simulated cattle was determined with two different ducting systems: a ram air inlet over the cab and NACA submerged inlets between the cab and trailer. The air flow within the conventional trailer was random and variable. The streamline vehicle with ram air inlet provided a nearly uniform air flow which could be controlled. The streamline vehicle with NACA submerged inlets provided better flow conditions than the conventional livestock trailer but not as uniform or controllable as the ram inlet configuration.

Effect of inlet disturbances on fan inlet noise during a static test

NASA Technical Reports Server (NTRS)

Bekofske, K. L.; Sheer, R. E., Jr.; Wang, J. C. F.

1977-01-01

Measurements of fan rotor inlet noise taken during static test situations are at variance with aircraft engine flight data. In particular, static tests generally yield a significantly higher tone at blade passage frequency than that measured during flight. To explain this discrepancy, the extent of the influence of inlet ground vortices and large-scale inlet turbulence on the forward-radiated fan noise measured at a static test facility was investigated. While such inlet disturbances were generated intentionally in an anechoic test chamber, far-field acoustic measurements and inlet flow-field hot-film mappings of a fan rotor were obtained. Experimental results indicate that the acoustic effect of such disturbances appears to be less severe for supersonic than for subsonic tip speeds. Further, a reverse flow that occurs on the exterior cowl in static test facilities appears to be an additional prime candidate for creating inlet disturbances and causing variance between flight and static acoustic data.

Experimental Investigation of a Large-Scale Low-Boom Inlet Concept

NASA Technical Reports Server (NTRS)

Hirt, Stefanie M.; Chima, Rodrick V.; Vyas, Manan A.; Wayman, Thomas R.; Conners, Timothy R.; Reger, Robert W.

2011-01-01

A large-scale low-boom inlet concept was tested in the NASA Glenn Research Center 8- x 6- foot Supersonic Wind Tunnel. The purpose of this test was to assess inlet performance, stability and operability at various Mach numbers and angles of attack. During this effort, two models were tested: a dual stream inlet designed to mimic potential aircraft flight hardware integrating a high-flow bypass stream; and a single stream inlet designed to study a configuration with a zero-degree external cowl angle and to permit surface visualization of the vortex generator flow on the internal centerbody surface. During the course of the test, the low-boom inlet concept was demonstrated to have high recovery, excellent buzz margin, and high operability. This paper will provide an overview of the setup, show a brief comparison of the dual stream and single stream inlet results, and examine the dual stream inlet characteristics.

Investigation of the effects of inlet shapes on fan noise radiation

NASA Technical Reports Server (NTRS)

Clark, T. L.; Slotboom, D. R.; Vaidya, P. G.

1981-01-01

The effect of inlet shape on forward radiated fan tone noise directivities was investigated under experimentally simplified zero flow conditions. Simulated fan tone noise was radiated to the far field through various shaped zero flow inlets. Baseline data were collected for the simplest baffled and unbaffled straight pipe inlets. These data compared well with prediction. The more general inlet shapes tested were the conical, circular, and exponential surfaces of revolution and an asymmetric inlet achieved by cutting a straight pipe inlet at an acute angle. Approximate theories were developed for these general shapes and some comparisons with data are presented. The conical and exponential shapes produced directivities that differed considerably from the baseline data while the circular shape produced directivities similar to the baseline data. The asymmetric inlet produced asymmetric directivities with significant reductions over the straight pipe data for some angles.

Numerical solution of a flow inside a labyrinth seal

NASA Astrophysics Data System (ADS)

Šimák, Jan; Straka, Petr; Pelant, Jaroslav

2012-04-01

The aim of this study is a behaviour of a flow inside a labyrinth seal on a rotating shaft. The labyrinth seal is a type of a non-contact seal where a leakage of a fluid is prevented by a rather complicated path, which the fluid has to overcome. In the presented case the sealed medium is the air and the seal is made by a system of 20 teeth on a rotating shaft situated against a smooth static surface. Centrifugal forces present due to the rotation of the shaft create vortices in each chamber and thus dissipate the axial velocity of the escaping air.The structure of the flow field inside the seal is studied through the use of numerical methods. Three-dimensional solution of the Navier-Stokes equations for turbulent flow is very time consuming. In order to reduce the computational time we can simplify our problem and solve it as an axisymmetric problem in a two-dimensional meridian plane. For this case we use a transformation of the Navier-Stokes equations and of the standard k-omega turbulence model into a cylindrical coordinate system. A finite volume method is used for the solution of the resulting problem. A one-side modification of the Riemann problem for boundary conditions is used at the inlet and at the outlet of the axisymmetric channel. The total pressure and total density (temperature) are to be used preferably at the inlet whereas the static pressure is used at the outlet for the compatibility. This idea yields physically relevant boundary conditions. The important characteristics such as a mass flow rate and a power loss, depending on a pressure ratio (1.1 - 4) and an angular velocity (1000 - 15000 rpm) are evaluated.

Calculation of two-dimensional inlet flow fields in a supersonic free stream: Program documentation and test cases

NASA Technical Reports Server (NTRS)

Biringen, S. H.; Mcmillan, O. J.

1980-01-01

The use of a computer code for the calculation of two dimensional inlet flow fields in a supersonic free stream and a nonorthogonal mesh-generation code are illustrated by specific examples. Input, output, and program operation and use are given and explained for the case of supercritical inlet operation at a subdesign Mach number (M Mach free stream = 2.09) for an isentropic-compression, drooped-cowl inlet. Source listings of the computer codes are also provided.

Investigation of Flow Separation in a Transonic-fan Linear Cascade Using Visualization Methods

NASA Technical Reports Server (NTRS)

Lepicovsky, Jan; Chima, Rodrick V.; Jett, Thomas A.; Bencic, Timothy J.; Weiland, Kenneth E.

2000-01-01

An extensive study into the nature of the separated flows on the suction side of modem transonic fan airfoils at high incidence is described in the paper. Suction surface.flow separation is an important flow characteristic that may significantly contribute to stall flutter in transonic fans. Flutter in axial turbomachines is a highly undesirable and dangerous self-excited mode of blade oscillations that can result in high cycle fatigue blade failure. The study basically focused on two visualization techniques: surface flow visualization using dye oils, and schlieren (and shadowgraph) flow visualization. The following key observations were made during the study. For subsonic inlet flow, the flow on the suction side of the blade is separated over a large portion of the blade, and the separated area increases with increasing inlet Mach number. For the supersonic inlet flow condition, the flow is attached from the leading edge up to the point where a bow shock from the upper neighboring blade hits the blade surface. Low cascade solidity, for the subsonic inlet flow, results in an increased area of separated flow. For supersonic flow conditions, a low solidity results in an improvement in flow over the suction surface. Finally, computational results modeling the transonic cascade flowfield illustrate our ability to simulate these flows numerically.

Pressure compensated flow control valve

DOEpatents

Minteer, Daniel J.

1999-01-01

The invention is an air flow control valve which is capable of maintaining a constant flow at the outlet despite changes in the inlet or outlet pressure. The device consists of a shell assembly with an inlet chamber and outlet chamber separated by a separation plate. The chambers are connected by an orifice. Also located within the inlet chamber is a port controller assembly. The port controller assembly consists of a differential pressure plate and port cap affixed thereon. The cap is able to slide in and out of the orifice separating the inlet and outlet chambers. When the pressure differential is sufficient, the differential pressure plate rises or falls to maintain a constant air flow. Movement of the port controller assembly does not require the use of seals, diaphragms, tight tolerances, bushings, bearings, hinges, guides, or lubricants.

The effect of inclusion of inlets in dual drainage modelling

NASA Astrophysics Data System (ADS)

Chang, Tsang-Jung; Wang, Chia-Ho; Chen, Albert S.; Djordjević, Slobodan

2018-04-01

In coupled sewer and surface flood modelling approaches, the flow process in gullies is often ignored although the overland flow is drained to sewer network via inlets and gullies. Therefore, the flow entering inlets is transferred to the sewer network immediately, which may lead to a different flood estimation than the reality. In this paper, we compared two modelling approach with and without considering the flow processes in gullies in the coupled sewer and surface modelling. Three historical flood events were adopted for model calibration and validation. The results showed that the inclusion of flow process in gullies can further improve the accuracy of urban flood modelling.

Numerical study of a scramjet engine flow field

NASA Technical Reports Server (NTRS)

Drummond, J. P.; Weidner, E. H.

1981-01-01

A computer program has been developed to analyze the turbulent reacting flow field in a two-dimensional scramjet engine configuration. The program numerically solves the full two-dimensional Navier-Stokes and species equations in the engine inlet and combustor, allowing consideration of flow separation and possible inlet-combustor interactions. The current work represents an intermediate step towards development of a three-dimensional program to analyze actual scramjet engine flow fields. Results from the current program are presented that predict the flow field for two inlet-combustor configurations, and comparisons of the program with experiment are given to allow assessment of the modeling that is employed.

Single-stage experimental evaluation of tandem-airfoil rotor and stator blading for compressors. Part 7: Data and performance for stage E

NASA Technical Reports Server (NTRS)

Cheatham, J. G.

1974-01-01

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.

Development of an Experimental Data Base to Validate Compressor-Face Boundary Conditions Used in Unsteady Inlet Flow Computations

NASA Technical Reports Server (NTRS)

Sajben, Miklos; Freund, Donald D.

1998-01-01

The ability to predict the dynamics of integrated inlet/compressor systems is an important part of designing high-speed propulsion systems. The boundaries of the performance envelope are often defined by undesirable transient phenomena in the inlet (unstart, buzz, etc.) in response to disturbances originated either in the engine or in the atmosphere. Stability margins used to compensate for the inability to accurately predict such processes lead to weight and performance penalties, which translate into a reduction in vehicle range. The prediction of transients in an inlet/compressor system requires either the coupling of two complex, unsteady codes (one for the inlet and one for the engine) or else a reliable characterization of the inlet/compressor interface, by specifying a boundary condition. In the context of engineering development programs, only the second option is viable economically. Computations of unsteady inlet flows invariably rely on simple compressor-face boundary conditions (CFBC's). Currently, customary conditions include choked flow, constant static pressure, constant axial velocity, constant Mach number or constant mass flow per unit area. These conditions are straightforward extensions of practices that are valid for and work well with steady inlet flows. Unfortunately, it is not at all likely that any flow property would stay constant during a complex system transient. At the start of this effort, no experimental observation existed that could be used to formulate of verify any of the CFBC'S. This lack of hard information represented a risk for a development program that has been recognized to be unacceptably large. The goal of the present effort was to generate such data. Disturbances reaching the compressor face in flight may have complex spatial structures and temporal histories. Small amplitude disturbances may be decomposed into acoustic, vorticity and entropy contributions that are uncoupled if the undisturbed flow is uniform. This study is focused on the response of an inlet/compressor system to acoustic disturbances. From the viewpoint of inlet computations, acoustic disturbances are clearly the most important, since they are the only ones capable of moving upstream. Convective and entropy disturbances may also produce upstream-moving acoustic waves, but such processes are outside the scope of the present study.

Low NOx nozzle tip for a pulverized solid fuel furnace

DOEpatents

Donais, Richard E; Hellewell, Todd D; Lewis, Robert D; Richards, Galen H; Towle, David P

2014-04-22

A nozzle tip [100] for a pulverized solid fuel pipe nozzle [200] of a pulverized solid fuel-fired furnace includes: a primary air shroud [120] having an inlet [102] and an outlet [104], wherein the inlet [102] receives a fuel flow [230]; and a flow splitter [180] disposed within the primary air shroud [120], wherein the flow splitter disperses particles in the fuel flow [230] to the outlet [104] to provide a fuel flow jet which reduces NOx in the pulverized solid fuel-fired furnace. In alternative embodiments, the flow splitter [180] may be wedge shaped and extend partially or entirely across the outlet [104]. In another alternative embodiment, flow splitter [180] may be moved forward toward the inlet [102] to create a recessed design.

A novel approach of magnetorheological abrasive fluid finishing with swirling-assisted inlet flow

NASA Astrophysics Data System (ADS)

Kheradmand, Saeid; Esmailian, Mojtaba; Fatahy, A.

Abrasive flow machining has been the pioneer of new finishing processes. Rotating workpiece and imposing a magnetic field using magnetorheological working medium are some assisting manipulations to improve surface finishing, because they can increase the forces on the workpiece surface. Similarly, swirling the inlet flow using stationary swirler vanes, as a novel idea, may also increase forces on the surface, and then raise the material removal, with a lower expense and energy consumption compared with the case of workpiece rotation. Thus, in this paper, surface roughness improvement is studied in a pipe with rotating inlet flow of a magnetorheological finishing medium under imposing a magnetic field. The results are compared with the case of rotating workpiece, using 3D numerical simulation. The governing hydrodynamic parameters are investigated in both cases to monitor the flow variations. It is shown that surface roughness is improved by rotating inlet flow. However, it is found that finishing in the entrance length of swirling-assisted inlet flow can be so economical for short length workpieces, compared with the case of rotating workpiece, with very near Ra values. By comparison of the numerical results and published experimental data, current study also shows the ability of the numerical simulation, as an inexpensive and efficient tool, to predict the surface roughness changes in finishing processes.

Literature search of publications concerning the prediction of dynamic inlet flow distortion and related topics

NASA Technical Reports Server (NTRS)

Schweikhhard, W. G.; Chen, Y. S.

1983-01-01

Publications prior to March 1981 were surveyed to determine inlet flow dynamic distortion prediction methods and to catalog experimental and analytical information concerning inlet flow dynamic distortion prediction methods and to catalog experimental and analytical information concerning inlet flow dynamics at the engine-inlet interface of conventional aircraft (excluding V/STOL). The sixty-five publications found are briefly summarized and tabulated according to topic and are cross-referenced according to content and nature of the investigation (e.g., predictive, experimental, analytical and types of tests). Three appendices include lists of references, authors, organizations and agencies conducting the studies. Also, selected materials summaries, introductions and conclusions - from the reports are included. Few reports were found covering methods for predicting the probable maximum distortion. The three predictive methods found are those of Melick, Jacox and Motycka. The latter two require extensive high response pressure measurements at the compressor face, while the Melick Technique can function with as few as one or two measurements.

Comprehensive theory of the Deans' switch as a variable flow splitter: fluid mechanics, mass balance, and system behavior.

PubMed

Boeker, Peter; Leppert, Jan; Mysliwietz, Bodo; Lammers, Peter Schulze

2013-10-01

The Deans' switch is an effluent switching device based on controlling flows of carrier gas instead of mechanical valves in the analytical flow path. This technique offers high inertness and a wear-free operation. Recently new monolithic microfluidic devices have become available. In these devices the whole flow system is integrated into a small metal device with low thermal mass and leak-tight connections. In contrast to a mechanical valve-based system, a flow-controlled system is more difficult to calculate. Usually the Deans' switch is used to switch one inlet to one of two outlets, by means of two auxiliary flows. However, the Deans' switch can also be used to deliver the GC effluent with a specific split ratio to both outlets. The calculation of the split ratio of the inlet flow to the two outlets is challenging because of the asymmetries of the flow resistances. This is especially the case, if one of the outlets is a vacuum device, such as a mass spectrometer, and the other an atmospheric detector, e.g. a flame ionization detector (FID) or an olfactory (sniffing) port. The capillary flows in gas chromatography are calculated with the Hagen-Poiseuille equation of the laminar, isothermal and compressible flow in circular tubes. The flow resistances in the new microfluidic devices have to be calculated with the corresponding equation for rectangular cross-section microchannels. The Hagen-Poiseuille equation underestimates the flow to a vacuum outlet. A corrected equation originating from the theory of rarefied flows is presented. The calculation of pressures and flows of a Deans' switch based chromatographic system is done by the solution of mass balances. A specific challenge is the consideration of the antidiffusion resistor between the two auxiliary gas lines of the Deans' switch. A full solution for the calculation of the Deans' switch including this restrictor is presented. Results from validation measurements are in good accordance with the developed theories. A spreadsheet-based flow calculator is part of the Supporting Information.

Computational fluid dynamics analysis in support of the simplex turbopump design

NASA Technical Reports Server (NTRS)

Garcia, Roberto; Griffin, Lisa W.; Benjamin, Theodore G.; Cornelison, Joni W.; Ruf, Joseph H.; Williams, Robert W.

1994-01-01

Simplex is a turbopump that is being developed at NASA/Marshall Space Flight Center (MSFC) by an in-house team. The turbopump consists of a single-stage centrifugal impeller, vaned-diffuser pump powered by a single-stage, axial, supersonic, partial admission turbine. The turbine is driven by warm gaseous oxygen tapped off of the hybrid motor to which it will be coupled. Rolling element bearings are cooled by the pumping fluid. Details of the configuration and operating conditions are given by Marsh. CFD has been used extensively to verify one-dimensional (1D) predictions, assess aerodynamic and hydrodynamic designs, and to provide flow environments. The complete primary flow path of the pump-end and the hot gas path of the turbine, excluding the inlet torus, have been analyzed. All CFD analyses conducted for the Simplex turbopump employed the pressure based Finite Difference Navier-Stokes (FDNS) code using a standard kappa-epsilon turbulence model with wall functions. More detailed results are presented by Garcia et. al. To support the team, loading and temperature results for the turbine rotor were provided as inputs to structural and thermal analyses, and blade loadings from the inducer were provided for structural analyses.

Quasi One-Dimensional Unsteady Modeling of External Compression Supersonic Inlets

NASA Technical Reports Server (NTRS)

Kopasakis, George; Connolly, Joseph W.; Kratz, Jonathan

2012-01-01

The AeroServoElasticity task under the NASA Supersonics Project is developing dynamic models of the propulsion system and the vehicle in order to conduct research for integrated vehicle dynamic performance. As part of this effort, a nonlinear quasi 1-dimensional model of an axisymmetric external compression supersonic inlet is being developed. The model utilizes compressible flow computational fluid dynamics to model the internal inlet segment as well as the external inlet portion between the cowl lip and normal shock, and compressible flow relations with flow propagation delay to model the oblique shocks upstream of the normal shock. The external compression portion between the cowl-lip and the normal shock is also modeled with leaking fluxes crossing the sonic boundary, with a moving CFD domain at the normal shock boundary. This model has been verified in steady state against tunnel inlet test data and it s a first attempt towards developing a more comprehensive model for inlet dynamics.

Analytical and experimental evaluation of a 3-D hypersonic fixed-geometry, swept, mixed compression inlet

NASA Technical Reports Server (NTRS)

Agnone, Anthony M.

1987-01-01

The performance of a fixed-geometry, swept, mixed compression hypersonic inlet is presented. The experimental evaluation was conducted for a Mach number of 6.0 and for several angles of attack. The measured surface pressures and pitot pressure surveys at the inlet throat are compared to computations using a three-dimensional Euler code and an integral boundary layer theory. Unique features of the intake design, including the boundary layer control, insure a high inlet performance. The experimental data show the inlet has a high mass averaged total pressure recovery, a high mass capture and nearly uniform flow diffusion. The swept inlet exhibits excellent starting characteristics, and high flow stability at angle of attack.

Design of a very-low-bleed Mach 2.5 mixed-compression inlet with 45 percent internal contraction

NASA Technical Reports Server (NTRS)

Wasserbauer, J. F.; Shaw, R. J.; Neumann, H. E.

1975-01-01

A full-scale, mixed-compression inlet was designed for operation with the TF30-P-3 turbofan engine and tested at Mach numbers of 2.5 and 2.0. The two-cone axisymmetric inlet had minimum internal contraction consistent with high total pressure recovery and low cowl drag. At Mach 2.5, inlet recovery was 0.906 with only 0.021 centerbody bleed mass-flow ratio and no cowl bleed. Increased centerbody bleed gave a maximum inlet unstart angle of attack of 6.85 deg. At Mach 2.0, inlet recovery was 0.94 with only 0.014 centerbody bleed mass-flow ratio and no cowl bleed. Inlet performance and angle-of-attack tolerance is presented for operation at Mach numbers of 2.5 and 2.0.

Inlet Flow Control and Prediction Technologies for Embedded Propulsion Systems

NASA Technical Reports Server (NTRS)

McMillan, Michelle L.; Gissen, Abe; Vukasinovic, Bojan; Lakebrink, Matthew T.; Glezer, Ari; Mani, Mori; Mace, James

2010-01-01

Fail-safe inlet flow control may enable high-speed cruise efficiency, low noise signature, and reduced fuel-burn goals for hybrid wing-body aircraft. The objectives of this program are to develop flow control and prediction methodologies for boundary-layer ingesting (BLI) inlets used in these aircraft. This report covers the second of a three year program. The approach integrates experiments and numerical simulations. Both passive and active flow-control devices were tested in a small-scale wind tunnel. Hybrid actuation approaches, combining a passive microvane and active synthetic jet, were tested in various geometric arrangements. Detailed flow measurements were taken to provide insight into the flow physics. Results of the numerical simulations were correlated against experimental data. The sensitivity of results to grid resolution and turbulence models was examined. Aerodynamic benefits from microvanes and microramps were assessed when installed in an offset BLI inlet. Benefits were quantified in terms of recovery and distortion changes. Microvanes were more effective than microramps at improving recovery and distortion.

Miniature piezo electric vacuum inlet valve

DOEpatents

Keville, Robert F.; Dietrich, Daniel D.

1998-03-24

A miniature piezo electric vacuum inlet valve having a fast pulse rate and is battery operated with variable flow capability. The low power (<1.6 watts), high pulse rate (<2 milliseconds), variable flow inlet valve is utilized for mass spectroscopic applications or other applications where pulsed or continuous flow conditions are needed. The inlet valve also has a very minimal dead volume of less than 0.01 std/cc. The valve can utilize, for example, a 12 Vdc input/750 Vdc, 3 mA output power supply compared to conventional piezo electric valves which require preloading of the crystal drive mechanism and 120 Vac, thus the valve of the present invention is smaller by a factor of three.

Analysis of the dynamic response of a supersonic inlet to flow-field perturbations upstream of the normal shock

NASA Technical Reports Server (NTRS)

Cole, G. L.; Willoh, R. G.

1975-01-01

A linearized mathematical analysis is presented for determining the response of normal shock position and subsonic duct pressures to flow-field perturbations upstream of the normal shock in mixed-compression supersonic inlets. The inlet duct cross-sectional area variation is approximated by constant-area sections; this approximation results in one-dimensional wave equations. A movable normal shock separates the supersonic and subsonic flow regions, and a choked exit is assumed for the inlet exit condition. The analysis leads to a closed-form matrix solution for the shock position and pressure transfer functions. Analytical frequency response results are compared with experimental data and a method of characteristics solution.

Aerodynamic design of gas and aerosol samplers for aircraft

NASA Technical Reports Server (NTRS)

Soderman, Paul T.; Hazen, Nathan L.; Brune, William H.

1991-01-01

The aerodynamic design of airborne probes for the capture of air and aerosols is discussed. Emphasis is placed on the key parameters that affect proper sampling, such as inlet-lip design, internal duct components for low pressure drop, and exhaust geometry. Inlet designs that avoid sonic flow conditions on the lip and flow separation in the duct are shown. Cross-stream velocities of aerosols are expressed in terms of droplet density and diameter. Flow curvature, which can cause aerosols to cross streamlines and impact on probe walls, can be minimized by means of a proper inlet shape and proper probe orientation, and by avoiding bends upstream of the test section. A NASA panel code called PMARC was used successfully to compute streamlines around aircraft and probes, as well as to compute to local velocity and pressure distributions in inlets. A NACA 1-series inlet with modified lip radius was used for the airborne capture of stratospheric chlorine monoxide at high altitude and high flight speed. The device has a two-stage inlet that decelerates the inflow with little disturbance to the flow through the test section. Diffuser design, exhaust hood design, valve loss, and corner vane geometry are discussed.

SUPIN: A Computational Tool for Supersonic Inlet Design

NASA Technical Reports Server (NTRS)

Slater, John W.

2016-01-01

A computational tool named SUPIN is being developed to design and analyze the aerodynamic performance of supersonic inlets. The inlet types available include the axisymmetric pitot, three-dimensional pitot, axisymmetric outward-turning, two-dimensional single-duct, two-dimensional bifurcated-duct, and streamline-traced inlets. The aerodynamic performance is characterized by the flow rates, total pressure recovery, and drag. The inlet flow-field is divided into parts to provide a framework for the geometry and aerodynamic modeling. Each part of the inlet is defined in terms of geometric factors. The low-fidelity aerodynamic analysis and design methods are based on analytic, empirical, and numerical methods which provide for quick design and analysis. SUPIN provides inlet geometry in the form of coordinates, surface angles, and cross-sectional areas. SUPIN can generate inlet surface grids and three-dimensional, structured volume grids for use with higher-fidelity computational fluid dynamics (CFD) analysis. Capabilities highlighted in this paper include the design and analysis of streamline-traced external-compression inlets, modeling of porous bleed, and the design and analysis of mixed-compression inlets. CFD analyses are used to verify the SUPIN results.

Establishing repeatable operation of a centrifugal compressor research facility for aerodynamic investigations

NASA Astrophysics Data System (ADS)

Dolan, Matthew Philip

The objective of this research has been to analyze the steady state performance of a new centrifugal compressor research facility. The CSTAR (Centrifugal STage for Aerodynamic Research) compressor has been designed to be placed as the last stage in an axial compressor and its performance in this flow regime was measured. Because the compressor was designed as a research vehicle, unique instrumentation throughout the flow path provides a detailed look at its steady state performance. Rakes at the inlet and deswirl section quantify the overall performance but other instrumentation is used to understand the component performance. Static pressure taps along the shroud, within the diffuser, and through the turn-to-axial show the static pressure rise. Additionally, rakes at the inlet and exit of diffuser also characterize the performance of the wedge diffuser and the impeller. Additionally, capacitance probes located at the knee and exducer of the impeller non-intrusively measure the size of the tip clearance during facility operation. An investigation into these measurements resulted in a standard procedure for in-situ calibration and installation to produce repeatable and accurate clearance measurements. Finally, the feasibility of future Laser Doppler Velocimetry measurements acquired through the shroud window was tested and was found to be achievable with the use of beam translators to ensure that measurement volumes are created after beam refraction through the windows. Inlet conditions of the facility have been investigated and fluctuations of the ambient conditions have been mitigated with a large settling chamber to ensure repeatable and stable operation. The current instrumentation was utilized to determine the compressor performance. Measurements of the steady performance parameters along with those of the internal flowfield are documented.

The Role of Design-of-Experiments in Managing Flow in Compact Air Vehicle Inlets

NASA Technical Reports Server (NTRS)

Anderson, Bernhard H.; Miller, Daniel N.; Gridley, Marvin C.; Agrell, Johan

2003-01-01

It is the purpose of this study to demonstrate the viability and economy of Design-of-Experiments methodologies to arrive at microscale secondary flow control array designs that maintain optimal inlet performance over a wide range of the mission variables and to explore how these statistical methods provide a better understanding of the management of flow in compact air vehicle inlets. These statistical design concepts were used to investigate the robustness properties of low unit strength micro-effector arrays. Low unit strength micro-effectors are micro-vanes set at very low angles-of-incidence with very long chord lengths. They were designed to influence the near wall inlet flow over an extended streamwise distance, and their advantage lies in low total pressure loss and high effectiveness in managing engine face distortion. The term robustness is used in this paper in the same sense as it is used in the industrial problem solving community. It refers to minimizing the effects of the hard-to-control factors that influence the development of a product or process. In Robustness Engineering, the effects of the hard-to-control factors are often called noise , and the hard-to-control factors themselves are referred to as the environmental variables or sometimes as the Taguchi noise variables. Hence Robust Optimization refers to minimizing the effects of the environmental or noise variables on the development (design) of a product or process. In the management of flow in compact inlets, the environmental or noise variables can be identified with the mission variables. Therefore this paper formulates a statistical design methodology that minimizes the impact of variations in the mission variables on inlet performance and demonstrates that these statistical design concepts can lead to simpler inlet flow management systems.

Design and aerodynamic performance evaluation of a high-work mixed flow turbine stage

NASA Technical Reports Server (NTRS)

Neri, Remo N.; Elliott, Thomas J.; Marsh, David N.; Civinskas, Kestutis C.

1994-01-01

As axial and radial turbine designs have been pushed to their aerothermodynamic and mechanical limits, the mixed-flow turbine (MFT) concept has been projected to offer performance and durability improvements, especially when ceramic materials are considered. The objective of this NASA/U.S. Army sponsored mixed-flow turbine (AMFT) program was to determine the level of performance attainable with MFT technology within the mechanical constraints of 1997 projected ceramic material properties. The MFT geometry is similar to a radial turbine, exhibiting a large radius change from inlet to exit, but differing in that the inlet flowpath is not purely radial, nor axial, but mixed; it is the inlet geometry that gives rise to the name 'mixed-flow'. The 'mixed' orientation of the turbine inlet offers several advantages over radial designs by allowing a nonzero inlet blade angle yet maintaining radial-element blades. The oblique inlet not only improves the particle-impact survivability of the design, but improves the aerodynamic performance by reducing the incidence at the blade inlet. The difficulty, however, of using mixed-flow geometry lies in the scarcity of detailed data and documented design experience. This paper reports the design of a MFT stage designed with the intent to maximize aerodynamic performance by optimizing design parameters such as stage reaction, rotor incidence, flowpath shape, blade shape, vane geometry, and airfoil counts using 2-D, 3-D inviscid, and 3-D viscous computational fluid dynamics code. The aerodynamic optimization was accomplished while maintaining mechanical integrity with respect to vibration and stress levels in the rotor. A full-scale cold-flow rig test was performed with metallic hardware fabricated to the specifications of the hot ceramic geometry to evaluate the stage performance.

Systems to facilitate reducing flashback/flame holding in combustion systems

DOEpatents

Lacy, Benjamin Paul [Greer, SC; Kraemer, Gilbert Otto [Greer, SC; Varatharajan, Balachandar [Clifton Park, NY; Yilmaz, Ertan [Albany, NY; Zuo, Baifang [Simpsonville, SC

2012-02-21

A method for assembling a premixing injector is provided. The method includes providing a centerbody including a center axis and a radially outer surface, and providing an inlet flow conditioner. The inlet flow conditioner includes a radially outer wall, a radially inner wall, and an end wall coupled substantially perpendicularly between the outer wall and the inner wall. Each of the outer wall and the end wall include a plurality of openings defined therein. The outer wall, the inner wall, and the end wall define a first passage therebetween. The method also includes coupling the inlet flow conditioner to the centerbody such that the inlet flow conditioner substantially circumscribes the centerbody, such that the inner wall is substantially parallel to the centerbody outer surface, and such that a second passage is defined between the centerbody outer surface and the inner wall.

Design and experimental verification of an equivalent forebody to produce disturbances equivalent to those of a forebody with flowing inlets

NASA Technical Reports Server (NTRS)

Haynes, Davy A.; Miller, David S.; Klein, John R.; Louie, Check M.

1988-01-01

A method by which a simple equivalent faired body can be designed to replace a more complex body with flowing inlets has been demonstrated for supersonic flow. An analytically defined, geometrically simple faired inlet forebody has been designed using a linear potential code to generate flow perturbations equivalent to those produced by a much more complex forebody with inlets. An equivalent forebody wind-tunnel model was fabricated and a test was conducted in NASA Langley Research Center's Unitary Plan Wind Tunnel. The test Mach number range was 1.60 to 2.16 for angles of attack of -4 to 16 deg. Test results indicate that, for the purposes considered here, the equivalent forebody simulates the original flowfield disturbances to an acceptable degree of accuracy.

A hybrid CFD/characteristics method for fast characterization of hypersonic blunt forebody/inlet flow

NASA Astrophysics Data System (ADS)

Gao, WenZhi; Li, ZhuFei; Yang, JiMing

2015-10-01

A hybrid CFD/characteristic method (CCM) was proposed for fast design and evaluation of hypersonic inlet flow with nose bluntness, which targets the combined advantages of CFD and method of characteristics. Both the accuracy and efficiency of the developed CCM were verified reliably, and it was well demonstrated for the external surfaces design of a hypersonic forebody/inlet with nose bluntness. With the help of CCM method, effects of nose bluntness on forebody shock shapes and the flowfield qualities which dominate inlet performance were examined and analyzed on the two-dimensional and axisymmetric configurations. The results showed that blunt effects of a wedge forebody are more substantial than that of related cone cases. For a conical forebody with a properly blunted nose, a recovery of the shock front back to that of corresponding sharp nose is exhibited, accompanied with a gradually fading out of entropy layer effects. Consequently a simplification is thought to be reasonable for an axisymmetric inlet with a proper compression angle, and a blunt nose of limited radius can be idealized as a sharp nose, as the spillage and flow variations at the entrance are negligible, even though the nose scale increases to 10% cowl lip radius. Whereas for two-dimensional inlets, the blunt effects are substantial since not only the inlet capturing/starting capabilities, but also the flow uniformities are obviously degraded.

The performance of a centrifugal compressor with high inlet prewhirl

DOE Office of Scientific and Technical Information (OSTI.GOV)

Whitfield, A.; Abdullah, A.H.

1998-07-01

The performance requirements of centrifugal compressors usually include a broad operating range between surge and choke. This becomes increasingly difficult to achieve as increased pressure ratio is demanded. In order to suppress the tendency to surge and extend the operating range at low flow rates, inlet swirl is often considered through the application of inlet guide vanes. To generate high inlet swirl angles efficiently, an inlet volute has been applied as the swirl generator, and a variable geometry design developed in order to provide zero swirl. The variable geometry approach can be applied to increase the swirl progressively or tomore » switch rapidly from zero swirl to maximum swirl. The variable geometry volute and the swirl conditions generated are described. The performance of a small centrifugal compressor is presented for a wide range of inlet swirl angles. In addition to the basic performance characteristics of the compressor, the onsets of flow reversals at impeller inlet are presented, together with the development of pressure pulsations, in the inlet and discharge ducts, through to full surge. The flow rate at which surge occurred was shown, by the shift of the peak pressure condition and by the measurement of the pressure pulsations, to be reduced by over 40%.« less

Adaptive compliant structures for flow regulation

PubMed Central

Brinkmeyer, Alex; Theunissen, Raf; M. Weaver, Paul; Pirrera, Alberto

2017-01-01

This paper introduces conceptual design principles for a novel class of adaptive structures that provide both flow regulation and control. While of general applicability, these design principles, which revolve around the idea of using the instabilities and elastically nonlinear behaviour of post-buckled panels, are exemplified through a case study: the design of a shape-adaptive air inlet. The inlet comprises a deformable post-buckled member that changes shape depending on the pressure field applied by the surrounding fluid, thereby regulating the inlet aperture. By tailoring the stress field in the post-buckled state and the geometry of the initial, stress-free configuration, the deformable section can snap through to close or open the inlet completely. Owing to its inherent ability to change shape in response to external stimuli—i.e. the aerodynamic loads imposed by different operating conditions—the inlet does not have to rely on linkages and mechanisms for actuation, unlike conventional flow-controlling devices. PMID:28878567

Adaptive compliant structures for flow regulation.

PubMed

Arena, Gaetano; M J Groh, Rainer; Brinkmeyer, Alex; Theunissen, Raf; M Weaver, Paul; Pirrera, Alberto

2017-08-01

This paper introduces conceptual design principles for a novel class of adaptive structures that provide both flow regulation and control. While of general applicability, these design principles, which revolve around the idea of using the instabilities and elastically nonlinear behaviour of post-buckled panels, are exemplified through a case study: the design of a shape-adaptive air inlet. The inlet comprises a deformable post-buckled member that changes shape depending on the pressure field applied by the surrounding fluid, thereby regulating the inlet aperture. By tailoring the stress field in the post-buckled state and the geometry of the initial, stress-free configuration, the deformable section can snap through to close or open the inlet completely. Owing to its inherent ability to change shape in response to external stimuli-i.e. the aerodynamic loads imposed by different operating conditions-the inlet does not have to rely on linkages and mechanisms for actuation, unlike conventional flow-controlling devices.

Monopolar fuel cell stack coupled together without use of top or bottom cover plates or tie rods

NASA Technical Reports Server (NTRS)

Narayanan, Sekharipuram R. (Inventor); Valdez, Thomas I. (Inventor)

2009-01-01

A monopolar fuel cell stack comprises a plurality of sealed unit cells coupled together. Each unit cell comprises two outer cathodes adjacent to corresponding membrane electrode assemblies and a center anode plate. An inlet and outlet manifold are coupled to the anode plate and communicate with a channel therein. Fuel flows from the inlet manifold through the channel in contact with the anode plate and flows out through the outlet manifold. The inlet and outlet manifolds are arranged to couple to the inlet and outlet manifolds respectively of an adjacent one of the plurality of unit cells to permit fuel flow in common into all of the inlet manifolds of the plurality of the unit cells when coupled together in a stack and out of all of the outlet manifolds of the plurality of unit cells when coupled together in a stack.

Analysis of DC control in double-inlet GM type pulse tube refrigerators for detectors

NASA Astrophysics Data System (ADS)

Du, B. Y.

2016-10-01

Pulse tube refrigerators have demonstrated many advantages with respect to temperature stability, vibration, reliability and lifetime among cryo-coolers for detectors. Double-inlet type pulse tube refrigerators are popular in GM type pulse tube refrigerators. The single double-inlet valve may introduce DC flow in refrigerator, which deteriorates the performance of pulse tube refrigerator. One new type of DC control mode is introduced in this paper. Two parallel-placed needle valves with opposite direction named double-valve configuration, instead of single double-inlet valve, are used in our experiment to reduce the DC flow. With two double-inlet operating, the lowest cold end temperature of 18.1K and a coolant of 1.2W@20K have been obtained. It has proved that this method is useful for controlling DC flow of the pulse tube refrigerators, which is very important to understand the characters of pulse tube refrigerators for detectors.

Investigation at supersonic and subsonic Mach numbers of auxiliary inlets supplying secondary air flow to ejector exhaust nozzles

NASA Technical Reports Server (NTRS)

Hearth, Donald P; Cubbison, Robert W

1956-01-01

The results indicated increases in auxiliary-inlet pressure recovery with increases in scoop height relative to the boundary-layer thickness. The pressure recovery increased at about the same rate as theoretically predicted for an inlet in a boundary layer having a one-seventh power profile, but was only about 0.68 to 0.75 of the theoretically obtainable values. Under some operating conditions, flow from the primary jet was exhausted through the auxiliary inlet. This phenomenon could be predicted from the ejector pumping characteristics.

Removable feedwater sparger assembly

DOEpatents

Challberg, R.C.

1994-10-04

A removable feedwater sparger assembly includes a sparger having an inlet pipe disposed in flow communication with the outlet end of a supply pipe. A tubular coupling includes an annular band fixedly joined to the sparger inlet pipe and a plurality of fingers extending from the band which are removably joined to a retention flange extending from the supply pipe for maintaining the sparger inlet pipe in flow communication with the supply pipe. The fingers are elastically deflectable for allowing engagement of the sparger inlet pipe with the supply pipe and for disengagement therewith. 8 figs.

Removable feedwater sparger assembly

DOEpatents

Challberg, Roy C.

1994-01-01

A removable feedwater sparger assembly includes a sparger having an inlet pipe disposed in flow communication with the outlet end of a supply pipe. A tubular coupling includes an annular band fixedly joined to the sparger inlet pipe and a plurality of fingers extending from the band which are removably joined to a retention flange extending from the supply pipe for maintaining the sparger inlet pipe in flow communication with the supply pipe. The fingers are elastically deflectable for allowing engagement of the sparger inlet pipe with the supply pipe and for disengagement therewith.

Investigation of X24C-2 10-Stage Axial-Flow Compressor. 2; Effect of Inlet-Air Pressure and Temperature of Performance

NASA Technical Reports Server (NTRS)

Finger, Harold B.; Schum, Harold J.; Buckner, Howard Jr.

1947-01-01

Effect of inlet-air pressure and temperature on the performance of the X24-2 10-Stage Axial-Flow Compressor from the X24C-2 turbojet engine was evaluated. Speeds of 80, 89, and 100 percent of equivalent design speed with inlet-air pressures of 6 and 12 inches of mercury absolute and inlet-air temperaures of approximately 538 degrees, 459 degrees,and 419 degrees R ( 79 degrees, 0 degrees, and minus 40 degrees F). Results were compared with prior investigations.

Flow-driven waves and sink-driven oscillations during aggregation of Dictyostelium discoideum

NASA Astrophysics Data System (ADS)

Gholami, Azam; Zykov, Vladimir; Steinbock, Oliver; Bodenschatz, Eberhard

The slime mold Dictyostelium discoideum (D.d) is a well-known model system for the study of biological pattern formation. Under starvation, D.d. cells aggregate chemotactically towards cAMP signals emitted periodically from an aggregation center. In the natural environment, D.d cells may experience fluid flows that can profoundly change the underlying wave generation process. We investigate spatial-temporal dynamics of a uniformly distributed population of D.d. cells in a flow-through narrow microfluidic channel with a cell-free inlet area. We show that flow can significantly influence the dynamics of the system and lead to a flow- driven instability that initiate downstream traveling cAMP waves. We also show that cell-free boundary regions have a significant effect on the observed patterns and can lead to a new kind of instability. Since there are no cells in the inlet to produce cAMP, the points in the vicinity of the inlet lose cAMP due to advection or diffusion and gain only a little from the upstream of the channel (inlet). In other words, there is a large negative flux of cAMP in the neighborhood close to the inlet, which can be considered as a sink. This negative flux close to the inlet drives a new kind of instability called sink-driven oscillations. Financial support of the MaxSynBio Consortium is acknowledged.

Validation of WIND for a Series of Inlet Flows

NASA Technical Reports Server (NTRS)

Slater, John W.; Abbott, John M.; Cavicchi, Richard H.

2002-01-01

Validation assessments compare WIND CFD simulations to experimental data for a series of inlet flows ranging in Mach number from low subsonic to hypersonic. The validation procedures follow the guidelines of the AIAA. The WIND code performs well in matching the available experimental data. The assessments demonstrate the use of WIND and provide confidence in its use for the analysis of aircraft inlets.

Streamline similarity method for flow distributions and shock losses at the impeller inlet of the centrifugal pump

NASA Astrophysics Data System (ADS)

Zhang, Zh.

2018-02-01

An analytical method is presented, which enables the non-uniform velocity and pressure distributions at the impeller inlet of a pump to be accurately computed. The analyses are based on the potential flow theory and the geometrical similarity of the streamline distribution along the leading edge of the impeller blades. The method is thus called streamline similarity method (SSM). The obtained geometrical form of the flow distribution is then simply described by the geometrical variable G( s) and the first structural constant G I . As clearly demonstrated and also validated by experiments, both the flow velocity and the pressure distributions at the impeller inlet are usually highly non-uniform. This knowledge is indispensible for impeller blade designs to fulfill the shockless inlet flow condition. By introducing the second structural constant G II , the paper also presents the simple and accurate computation of the shock loss, which occurs at the impeller inlet. The introduction of two structural constants contributes immensely to the enhancement of the computational accuracies. As further indicated, all computations presented in this paper can also be well applied to the non-uniform exit flow out of an impeller of the Francis turbine for accurately computing the related mean values.

Numerical simulation of transonic compressor under circumferential inlet distortion and rotor/stator interference using harmonic balance method

NASA Astrophysics Data System (ADS)

Wang, Ziwei; Jiang, Xiong; Chen, Ti; Hao, Yan; Qiu, Min

2018-05-01

Simulating the unsteady flow of compressor under circumferential inlet distortion and rotor/stator interference would need full-annulus grid with a dual time method. This process is time consuming and needs a large amount of computational resources. Harmonic balance method simulates the unsteady flow in compressor on single passage grid with a series of steady simulations. This will largely increase the computational efficiency in comparison with the dual time method. However, most simulations with harmonic balance method are conducted on the flow under either circumferential inlet distortion or rotor/stator interference. Based on an in-house CFD code, the harmonic balance method is applied in the simulation of flow in the NASA Stage 35 under both circumferential inlet distortion and rotor/stator interference. As the unsteady flow is influenced by two different unsteady disturbances, it leads to the computational instability. The instability can be avoided by coupling the harmonic balance method with an optimizing algorithm. The computational result of harmonic balance method is compared with the result of full-annulus simulation. It denotes that, the harmonic balance method simulates the flow under circumferential inlet distortion and rotor/stator interference as precise as the full-annulus simulation with a speed-up of about 8 times.

Entropy considerations applied to shock unsteadiness in hypersonic inlets

NASA Astrophysics Data System (ADS)

Bussey, Gillian Mary Harding

The stability of curved or rectangular shocks in hypersonic inlets in response to flow perturbations can be determined analytically from the principle of minimum entropy. Unsteady shock wave motion can have a significant effect on the flow in a hypersonic inlet or combustor. According to the principle of minimum entropy, a stable thermodynamic state is one with the lowest entropy gain. A model based on piston theory and its limits has been developed for applying the principle of minimum entropy to quasi-steady flow. Relations are derived for analyzing the time-averaged entropy gain flux across a shock for quasi-steady perturbations in atmospheric conditions and angle as a perturbation in entropy gain flux from the steady state. Initial results from sweeping a wedge at Mach 10 through several degrees in AEDC's Tunnel 9 indicates the bow shock becomes unsteady near the predicted normal Mach number. Several curved shocks of varying curvature are compared to a straight shock with the same mean normal Mach number, pressure ratio, or temperature ratio. The present work provides analysis and guidelines for designing an inlet robust to off- design flight or perturbations in flow conditions an inlet is likely to face. It also suggests that inlets with curved shocks are less robust to off-design flight than those with straight shocks such as rectangular inlets. Relations for evaluating entropy perturbations for highly unsteady flow across a shock and limits on their use were also developed. The normal Mach number at which a shock could be stable to high frequency upstream perturbations increases as the speed of the shock motion increases and slightly decreases as the perturbation size increases. The present work advances the principle of minimum entropy theory by providing additional validity for using the theory for time-varying flows and applying it to shocks, specifically those in inlets. While this analytic tool is applied in the present work for evaluating the stability of shocks in hypersonic inlets, it can be used for an arbitrary application with a shock.

Effects of inlet boundary conditions, on the computed flow in the Turbine-99 draft tube, using OpenFOAM and CFX

NASA Astrophysics Data System (ADS)

Nilsson, H.; Cervantes, M. J.

2012-11-01

The flow in the Turbine-99 Kaplan draft tube was thoroughly investigated at three workshops (1999, 2001, 2005), which aimed at determining the state of the art of draft tube simulations. The flow is challenging due to the different flow phenomena appearing simultaneously such as unsteadiness, separation, swirl, turbulence, and a strong adverse pressure gradient. The geometry and the experimentally determined inlet boundary conditions were provided to the Turbine-99 workshop participants. At the final workshop, angular resolved inlet velocity boundary conditions were provided. The rotating non-axi-symmetry of the inlet flow due to the runner blades was thus included. The effect of the rotating angular resolution was however not fully investigated at that workshop. The first purpose of this work is to further investigate this effect. Several different inlet boundary conditions are applied - the angular resolved experimental data distributed at the Turbine-99 workshop, the angular resolved results of a runner simulation with interpolated values using different resolution in the tangential and radial directions, and an axi-symmetric variant of the same numerical data. The second purpose of this work is to compare the results from the OpenFOAM and CFX CFD codes, using as similar settings as possible. The present results suggest that the experimental angular inlet boundary conditions proposed to the workshop are not adequate to simulate accurately the flow in the T-99 draft tube. The reason for this is that the experimental phase-averaged data has some important differences compared to the previously measured time-averaged data. Using the interpolated data from the runner simulation as inlet boundary condition however gives good results as long as the resolution of that data is sufficient. It is shown that the difference between the results using the angular-resolved and the corresponding symmetric inlet data is very small, suggesting that the importance of the angular resolution is small. The results from OpenFOAM and CFX are very similar as long as the inlet data resolution is fine enough. CFX seems to be more sensitive to that resolution.

Suction performance and internal flow of a 2-bladed helical inducer with inlet asymmetric plate

NASA Astrophysics Data System (ADS)

Watanabe, S.; Uchinono, Y.; Ishizaka, K.; Furukawa, A.; Kim, J.-H.

2013-10-01

It has been found in our past studies that the installation of asymmetric plate at the inlet of inducer is effective for the suppression of cavitation surge phenomenon. In the present study, the suction performance of 2-bladed helical inducer with an inlet asymmetric plate is experimentally investigated. It is observed that the suction performance in large flow rate conditions is not significantly influenced by the asymmetric plate, whereas the head of inducer with the asymmetric plate increases just before the head breakdown in partial flow conditions. To understand the mechanism of this additional head, the flow measurements and the numerical simulations are carried out. It is found that the circumferential component of absolute velocity at the exit of inducer slightly increases with the development of cavitation in both cases with and without the inlet asymmetric plate, indicating the increase of the theoretical head. The theoretical head increase with the inlet asymmetric plate is also confirmed by the unsteady numerical simulations, suggesting that the additional head is achieved through the increase of the theoretical head with the change of the exiting flow from the inducer associated with some amount of cavitation.

An investigation of several NACA 1 series axisymmetric inlets at Mach numbers from 0.4 to 1.29. [wind tunnel tests over range of mass-flow ratios and at angle of attack

NASA Technical Reports Server (NTRS)

Re, R. J.

1974-01-01

An investigation was conducted in the Langley 16-foot transonic tunnel to determine the performance of seven inlets having NACA 1-series contours and one inlet having an elliptical contour over a range of mass-flow ratios and at angle of attack. The inlet diameter ratio varied from 0.81 to 0.89; inlet length ratio varied from 0.75 to 1.25; and internal contraction ratio varied from 1.009 to 1.093. Reynolds number based on inlet maximum diameter varied from 3.4 million at a Mach number of 0.4 to 5.6 million at a Mach number of 1.29.

On the inlet vortex system. [preventing jet engine damage caused by debris pick-up

NASA Technical Reports Server (NTRS)

Bissinger, N. C.; Braun, G. W.

1974-01-01

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.

Experimental Evaluation of the Effect of Angle-of-attack on the External Aerodynamics and Mass Capture of a Symmetric Three-engine Air-breathing Launch Vehicle Configuration at Supersonic Speeds

NASA Technical Reports Server (NTRS)

Kim, Hyun D.; Frate, Franco C.

2001-01-01

A subscale aerodynamic model of the GTX air-breathing launch vehicle was tested at NASA Glenn Research Center's 10- by 10-Foot Supersonic Wind Tunnel from Mach 2.0 to 3.5 at various angles-of-attack. The objective of the test was to investigate the effect of angle-of-attack on inlet mass capture, inlet diverter effectiveness, and the flowfield at the cowl lip plane. The flow-through inlets were tested with and without boundary-layer diverters. Quantitative measurements such as inlet mass flow rates and pitot-pressure distributions in the cowl lip plane are presented. At a 3deg angle-of-attack, the flow rates for the top and side inlets were within 8 percent of the zero angle-of-attack value, and little distortion was evident at the cowl lip plane. Surface oil flow patterns showing the shock/boundary-layer interaction caused by the inlet spikes are shown. In addition to inlet data, vehicle forebody static pressure distributions, boundary-layer profiles, and temperature-sensitive paint images to evaluate the boundary-layer transition are presented. Three-dimensional parabolized Navier-Stokes computational fluid dynamics calculations of the forebody flowfield are presented and show good agreement with the experimental static pressure distributions and boundary-layer profiles. With the boundary-layer diverters installed, no adverse aerodynamic phenomena were found that would prevent the inlets from operating at the required angles-of-attack. We recommend that phase 2 of the test program be initiated, where inlet contraction ratio and diverter geometry variations will be tested.

The 3D Navier-Stokes analysis of a Mach 2.68 bifurcated rectangular mixed-compression inlet

NASA Technical Reports Server (NTRS)

Mizukami, M.; Saunders, J. D.

1995-01-01

The supersonic diffuser of a Mach 2.68 bifurcated, rectangular, mixed-compression inlet was analyzed using a three-dimensional (3D) Navier-Stokes flow solver. A two-equation turbulence model, and a porous bleed model based on unchoked bleed hole discharge coefficients were used. Comparisons were made with experimental data, inviscid theory, and two-dimensional Navier-Stokes analyses. The main objective was to gain insight into the inlet fluid dynamics. Examination of the computational results along with the experimental data suggest that the cowl shock-sidewall boundary layer interaction near the leading edge caused a substantial separation in the wind tunnel inlet model. As a result, the inlet performance may have been compromised by increased spillage and higher bleed mass flow requirements. The internal flow contained substantial waves that were not in the original inviscid design. 3D effects were fairly minor for this inlet at on-design conditions. Navier-Stokes analysis appears to be an useful tool for gaining insight into the inlet fluid dynamics. It provides a higher fidelity simulation of the flowfield than the original inviscid design, by taking into account boundary layers, porous bleed, and their interactions with shock waves.

Study of Gas Solid Flow Characteristics in Cyclone Inlet Ducts of A300Mwe CFB Boiler

NASA Astrophysics Data System (ADS)

Tang, J. Y.; Lu, X. F.; Lai, J.; Liu, H. Z.

Gas solid flow characteristics in cyclone's inlet duct of a 300MW CFB boiler were studied in a cold circulating fluidized bed (CFB) experimental setup according to a 410t/h CFB boiler with a scale of 10∶1. Tracer particles were adopted in the experiment and their motion trajectories in the two kinds of cyclone's inlet ducts were photographed by a high-speed camera. By analyzing the motion trajectories of tracer particles, acceleration performance of particle phases in the two inlet ducts was obtained. Results indicate that the acceleration performance of particles in the long inlet duct is better than that in the short inlet duct, but the pressure drop of the long inlet duct is higher. Meanwhile, under the same operating conditions, both the separation efficiency and the pressure drop of the cyclone are higher when the cyclone is connected with the long inlet duct. Figs 11, Tabs 4 and refs 10.

Active microchannel fluid processing unit and method of making

DOEpatents

Bennett, Wendy D [Kennewick, WA; Martin, Peter M [Kennewick, WA; Matson, Dean W [Kennewick, WA; Roberts, Gary L [West Richland, WA; Stewart, Donald C [Richland, WA; Tonkovich, Annalee Y [Pasco, WA; Zilka, Jennifer L [Pasco, WA; Schmitt, Stephen C [Dublin, OH; Werner, Timothy M [Columbus, OH

2001-01-01

The present invention is an active microchannel fluid processing unit and method of making, both relying on having (a) at least one inner thin sheet; (b) at least one outer thin sheet; (c) defining at least one first sub-assembly for performing at least one first unit operation by stacking a first of the at least one inner thin sheet in alternating contact with a first of the at least one outer thin sheet into a first stack and placing an end block on the at least one inner thin sheet, the at least one first sub-assembly having at least a first inlet and a first outlet; and (d) defining at least one second sub-assembly for performing at least one second unit operation either as a second flow path within the first stack or by stacking a second of the at least one inner thin sheet in alternating contact with second of the at least one outer thin sheet as a second stack, the at least one second sub-assembly having at least a second inlet and a second outlet.

Active microchannel fluid processing unit and method of making

DOEpatents

Bennett, Wendy D [Kennewick, WA; Martin, Peter M [Kennewick, WA; Matson, Dean W [Kennewick, WA; Roberts, Gary L [West Richland, WA; Stewart, Donald C [Richland, WA; Tonkovich, Annalee Y [Pasco, WA; Zilka, Jennifer L [Pasco, WA; Schmitt, Stephen C [Dublin, OH; Werner, Timothy M [Columbus, OH

2002-12-10

The present invention is an active microchannel fluid processing unit and method of making, both relying on having (a) at least one inner thin sheet; (b) at least one outer thin sheet; (c) defining at least one first sub-assembly for performing at least one first unit operation by stacking a first of the at least one inner thin sheet in alternating contact with a first of the at least one outer thin sheet into a first stack and placing an end block on the at least one inner thin sheet, the at least one first sub-assembly having at least a first inlet and a first outlet; and (d) defining at least one second sub-assembly for performing at least one second unit operation either as a second flow path within the first stack or by stacking a second of the at least one inner thin sheet in alternating contact with second of the at least one outer thin sheet as a second stack, the at least one second sub-assembly having at least a second inlet and a second outlet.

Flow Through a Rectangular-to-Semiannular Diffusing Transition Duct

NASA Technical Reports Server (NTRS)

Foster, Jeff; Wendt, Bruce J.; Reichert, Bruce A.; Okiishi, Theodore H.

1997-01-01

Rectangular-to-semiannular diffusing transition ducts are critical inlet components on supersonic airplanes having bifucated engine inlets. This paper documents measured details of the flow through a rectangular-to-semiannular transition duct having an expansion area ratio of 1.53. Three-dimensional velocity vectors and total pressures at the exit plane of the diffuser are presented. Surface oil-flow visualization and surface static pressure data are shown. The tests were conducted with an inlet Mach number of 0.786 and a Reynolds number based on the inlet centerline velocity and exit diameter of 3.2 x 10(exp 6). The measured data are compared with previously published computational results. The ability of vortex generators to reduce circumferential total pressure distortion is demonstrated.

Miniature piezo electric vacuum inlet valve

DOEpatents

Keville, R.F.; Dietrich, D.D.

1998-03-24

A miniature piezo electric vacuum inlet valve having a fast pulse rate and is battery operated with variable flow capability is disclosed. The low power (<1.6 watts), high pulse rate (<2 milliseconds), variable flow inlet valve is utilized for mass spectroscopic applications or other applications where pulsed or continuous flow conditions are needed. The inlet valve also has a very minimal dead volume of less than 0.01 std/cc. The valve can utilize, for example, a 12 Vdc input/750 Vdc, 3 mA output power supply compared to conventional piezo electric valves which require preloading of the crystal drive mechanism and 120 Vac, thus the valve of the present invention is smaller by a factor of three. 6 figs.

Large perturbation flow field analysis and simulation for supersonic inlets

NASA Technical Reports Server (NTRS)

Varner, M. O.; Martindale, W. R.; Phares, W. J.; Kneile, K. R.; Adams, J. C., Jr.

1984-01-01

An analysis technique for simulation of supersonic mixed compression inlets with large flow field perturbations is presented. The approach is based upon a quasi-one-dimensional inviscid unsteady formulation which includes engineering models of unstart/restart, bleed, bypass, and geometry effects. Numerical solution of the governing time dependent equations of motion is accomplished through a shock capturing finite difference algorithm, of which five separate approaches are evaluated. Comparison with experimental supersonic wind tunnel data is presented to verify the present approach for a wide range of transient inlet flow conditions.

Computer programs for calculating potential flow in propulsion system inlets

NASA Technical Reports Server (NTRS)

Stockman, N. O.; Button, S. L.

1973-01-01

In the course of designing inlets, particularly for VTOL and STOL propulsion systems, a calculational procedure utilizing three computer programs evolved. The chief program is the Douglas axisymmetric potential flow program called EOD which calculates the incompressible potential flow about arbitrary axisymmetric bodies. The other two programs, original with Lewis, are called SCIRCL AND COMBYN. Program SCIRCL generates input for EOD from various specified analytic shapes for the inlet components. Program COMBYN takes basic solutions output by EOD and combines them into solutions of interest, and applies a compressibility correction.

Validation of a three-dimensional viscous analysis of axisymmetric supersonic inlet flow fields

NASA Technical Reports Server (NTRS)

Benson, T. J.; Anderson, B. H.

1983-01-01

A three-dimensional viscous marching analysis for supersonic inlets was developed. To verify this analysis several benchmark axisymmetric test configurations were studied and are compared to experimental data. Detailed two-dimensional results for shock-boundary layer interactions are presented for flows with and without boundary layer bleed. Three dimensional calculations of a cone at angle of attack and a full inlet at attack are also discussed and evaluated. Results of the calculations demonstrate the code's ability to predict complex flow fields and establish guidelines for future calculations using similar codes.

Low pressure gas flow analysis through an effusive inlet using mass spectrometry

NASA Technical Reports Server (NTRS)

Brown, David R.; Brown, Kenneth G.

1988-01-01

A mass spectrometric method for analyzing flow past and through an effusive inlet designed for use on the tethered satellite and other entering vehicles is discussed. Source stream concentrations of species in a gaseous mixture are determined using a calibration of measured mass spectral intensities versus source stream pressure for standard gas mixtures and pure gases. Concentrations are shown to be accurate within experimental error. Theoretical explanations for observed mass discrimination effects as they relate to the various flow situations in the effusive inlet and the experimental apparatus are discussed.

Axial-Flow Turbine Rotor Discharge-Flow Overexpansion and Limit-Loading Condition, Part I: Computational Fluid Dynamics (CFD) Investigation

NASA Technical Reports Server (NTRS)

Chen, Shu-Cheng S.

2017-01-01

A Computational Fluid Dynamic (CFD) investigation is conducted over a two-dimensional axial-flow turbine rotor blade row to study the phenomena of turbine rotor discharge flow overexpansion at subcritical, critical, and supercritical conditions. Quantitative data of the mean-flow Mach numbers, mean-flow angles, the tangential blade pressure forces, the mean-flow mass flux, and the flow-path total pressure loss coefficients, averaged or integrated across the two-dimensional computational domain encompassing two blade-passages, are obtained over a series of 14 inlet-total to exit-static pressure ratios, from 1.5 (un-choked; subcritical condition) to 10.0 (supercritical with excessively high pressure ratio.) Detailed flow features over the full domain-of-computation, such as the streamline patterns, Mach contours, pressure contours, blade surface pressure distributions, etc. are collected and displayed in this paper. A formal, quantitative definition of the limit loading condition based on the channel flow theory is proposed and explained. Contrary to the comments made in the historical works performed on this subject, about the deficiency of the theoretical methods applied in analyzing this phenomena, using modern CFD method for the study of this subject appears to be quite adequate and successful. This paper describes the CFD work and its findings.

Model Scramjet Inlet Unstart Induced by Mass Addition and Heat Release

NASA Astrophysics Data System (ADS)

Im, Seong-Kyun; Baccarella, Damiano; McGann, Brendan; Liu, Qili; Wermer, Lydiy; Do, Hyungrok

2015-11-01

The inlet unstart phenomena in a model scramjet are investigated at an arc-heated hypersonic wind tunnel. The unstart induced by nitrogen or ethylene jets at low or high enthalpy Mach 4.5 freestream flow conditions are compared. The jet injection pressurizes the downstream flow by mass addition and flow blockage. In case of the ethylene jet injection, heat release from combustion increases the backpressure further. Time-resolved schlieren imaging is performed at the jet and the lip of the model inlet to visualize the flow features during unstart. High frequency pressure measurements are used to provide information on pressure fluctuation at the scramjet wall. In both of the mass and heat release driven unstart cases, it is observed that there are similar flow transient and quasi-steady behaviors of unstart shockwave system during the unstart processes. Combustion driven unstart induces severe oscillatory flow motions of the jet and the unstart shock at the lip of the scramjet inlet after the completion of the unstart process, while the unstarted flow induced by solely mass addition remains relatively steady. The discrepancies between the processes of mass and heat release driven unstart are explained by flow choking mechanism.

Investigation on inlet recirculation characteristics of double suction centrifugal compressor with unsymmetrical inlet

NASA Astrophysics Data System (ADS)

Yang, Ce; Wang, Yingjun; Lao, Dazhong; Tong, Ding; Wei, Longyu; Liu, Yixiong

2016-08-01

The inlet recirculation characteristics of double suction centrifugal compressor with unsymmetrical inlet structures were studied in numerical method, mainly focused on three issues including the amounts and differences of the inlet recirculation in different working conditions, the circumferential non-uniform distributions of the inlet recirculation, the recirculation velocity distributions of the upstream slot of the rear impeller. The results show that there are some differences between the recirculation of the front impeller and that of the rear impeller in whole working conditions. In design speed, the recirculation flow rate of the rear impeller is larger than that of the front impeller in the large flow range, but in the small flow range, the recirculation flow rate of the rear impeller is smaller than that of the front impeller. In different working conditions, the recirculation velocity distributions of the front and rear impeller are non-uniform along the circumferential direction and their non-uniform extents are quite different. The circumferential non-uniform extent of the recirculation velocity varies with the working conditions change. The circumferential non-uniform extent of the recirculation velocity of front impeller and its distribution are determined by the static pressure distribution of the front impeller, but that of the rear impeller is decided by the coupling effects of the inlet flow distortion of the rear impeller, the circumferential unsymmetrical distribution of the upstream slot and the asymmetric structure of the volute. In the design flow and small flow conditions, the recirculation velocities at different circumferential positions of the mean line of the upstream slot cross-section of the rear impeller are quite different, and the recirculation velocities distribution forms at both sides of the mean line are different. The recirculation velocity distributions in the cross-section of the upstream slot depend on the static pressure distributions in the intake duct.

Numerical simulation of calcium sulfate (CaSO4) fouling in the plate heat exchanger

NASA Astrophysics Data System (ADS)

Xu, Zhiming; Zhao, Yu; Han, Zhimin; Wang, Jingtao

2018-07-01

Plate heat exchanger is a widely used apparatus in the industrial production processes. Through a numerical simulation method, this paper calculates the deposition rate of CaSO4 fouling on heat transfer surfaces of the plate heat exchanger under saturation in the bulk. The effects of CaSO4 concentration in the range 0.7 kg/m3 to 1.5 kg/m3, inlet flow velocity under turbulent flow, and the fluid's inlet temperature from 288 K to 328 K on the deposition rate, removal mass rate and fouling resistance are investigated. The simulation results are compared with the experimental results showing similar trend. The simulation results show that the concentration and the flow velocity affect significantly the fouling characteristics in the plate heat exchanger. The deposition mass rate, removal mass rate, and asymptotic value of fouling resistance all increase with the increase in CaSO4 concentration and the inlet temperature of the hot fluid, while the asymptotic value of fouling resistance decreases with the increasing of inlet flow velocity. The influence of the inlet temperature of cold fluid may be negligible.

High Reynolds Number Investigation of a Flush-Mounted, S-Duct Inlet With Large Amounts of Boundary Layer Ingestion

NASA Technical Reports Server (NTRS)

Berrier, Bobby L.; Carter, Melissa B.; Allan, Brian G.

2005-01-01

An experimental investigation of a flush-mounted, S-duct inlet with large amounts of boundary layer ingestion has been conducted at Reynolds numbers up to full scale. The study was conducted in the NASA Langley Research Center 0.3-Meter Transonic Cryogenic Tunnel. In addition, a supplemental computational study on one of the inlet configurations was conducted using the Navier-Stokes flow solver, OVERFLOW. Tests were conducted at Mach numbers from 0.25 to 0.83, Reynolds numbers (based on aerodynamic interface plane diameter) from 5.1 million to 13.9 million (full-scale value), and inlet mass-flow ratios from 0.29 to 1.22, depending on Mach number. Results of the study indicated that increasing Mach number, increasing boundary layer thickness (relative to inlet height) or ingesting a boundary layer with a distorted profile decreased inlet performance. At Mach numbers above 0.4, increasing inlet airflow increased inlet pressure recovery but also increased distortion. Finally, inlet distortion was found to be relatively insensitive to Reynolds number, but pressure recovery increased slightly with increasing Reynolds number.

Assessing Fan Flutter Stability in the Presence of Inlet Distortion Using One-way and Two-way Coupled Methods

NASA Technical Reports Server (NTRS)

Herrick, Gregory P.

2014-01-01

Concerns regarding noise, propulsive efficiency, and fuel burn are inspiring aircraft designs wherein the propulsive turbomachines are partially (or fully)embedded within the airframe; such designs present serious concerns with regard to aerodynamic and aeromechanic performance of the compression system in response to inlet distortion. Previously, a preliminary design of a forward-swept high-speed fan exhibited flutter concerns in clean-inlet flows, and the present author then studied this fan further in the presence of off-design distorted in-flows. A three-dimensional, unsteady, Navier-Stokes computational fluid dynamics code is applied to analyze and corroborate fan performance with clean inlet flow. This code, already validated in its application to assess aerodynamic damping of vibrating blades at various flow conditions using a loosely-coupled approach, is modified to include a tightly-coupled aeroelastic simulation capability, and then loosely-coupled and tightly-coupled methods arecompared in their evaluation of flutter stability in distorted in-flows.

Estimation of inlet flow rates for image-based aneurysm CFD models: where and how to begin?

PubMed

Valen-Sendstad, Kristian; Piccinelli, Marina; KrishnankuttyRema, Resmi; Steinman, David A

2015-06-01

Patient-specific flow rates are rarely available for image-based computational fluid dynamics models. Instead, flow rates are often assumed to scale according to the diameters of the arteries of interest. Our goal was to determine how choice of inlet location and scaling law affect such model-based estimation of inflow rates. We focused on 37 internal carotid artery (ICA) aneurysm cases from the Aneurisk cohort. An average ICA flow rate of 245 mL min(-1) was assumed from the literature, and then rescaled for each case according to its inlet diameter squared (assuming a fixed velocity) or cubed (assuming a fixed wall shear stress). Scaling was based on diameters measured at various consistent anatomical locations along the models. Choice of location introduced a modest 17% average uncertainty in model-based flow rate, but within individual cases estimated flow rates could vary by >100 mL min(-1). A square law was found to be more consistent with physiological flow rates than a cube law. Although impact of parent artery truncation on downstream flow patterns is well studied, our study highlights a more insidious and potentially equal impact of truncation site and scaling law on the uncertainty of assumed inlet flow rates and thus, potentially, downstream flow patterns.

Indirect evaporative cooler using membrane-contained, liquid desiccant for dehumidification

DOEpatents

Kozubal, Eric Joseph

2016-12-13

An indirect evaporative cooler for cooling inlet supply air from a first temperature to a second, lower temperature using a stream of liquid coolant and a stream of exhaust or purge air. The cooler includes a first flow channel for inlet supply air and a second flow channel adjacent the first for exhaust air. The first and second flow channels are defined in part by sheets of a membrane permeable to water vapor such that mass is transferred as a vapor through the membrane from the inlet supply air to a contained liquid desiccant for dehumidification and also to the exhaust air as heat is transferred from the inlet supply air to the liquid coolant. A separation wall divides the liquid desiccant and the coolant but allows heat to be transferred from the supply air to the coolant which releases water vapor to the counter or cross flowing exhaust air.

Indirect evaporative cooler using membrane-contained, liquid desiccant for dehumidification

DOEpatents

Kozubal, Eric Joseph; Slayzak, Steven Joseph

2014-07-08

An indirect evaporative cooler for cooling inlet supply air from a first temperature to a second, lower temperature using a stream of liquid coolant and a stream of exhaust or purge air. The cooler includes a first flow channel for inlet supply air and a second flow channel adjacent the first for exhaust air. The first and second flow channels are defined in part by sheets of a membrane permeable to water vapor such that mass is transferred as a vapor through the membrane from the inlet supply air to a contained liquid desiccant for dehumidification and also to the exhaust air as heat is transferred from the inlet supply air to the liquid coolant. A separation wall divides the liquid desiccant and the coolant but allows heat to be transferred from the supply air to the coolant which releases water vapor to the counter or cross flowing exhaust air.

Cooled High-temperature Radial Turbine Program 2

NASA Technical Reports Server (NTRS)

Snyder, Philip H.

1991-01-01

The objective of this program was the design and fabrication of a air-cooled high-temperature radial turbine (HTRT) intended for experimental evaluation in a warm turbine test facility at the LeRC. The rotor and vane were designed to be tested as a scaled version (rotor diameter of 14.4 inches diameter) of a 8.021 inch diameter rotor designed to be capable of operating with a rotor inlet temperature (RIT) of 2300 F, a nominal mass flow of 4.56 lbm/sec, a work level of equal or greater than 187 Btu/lbm, and efficiency of 86 percent or greater. The rotor was also evaluated to determine it's feasibility to operate at 2500 F RIT. The rotor design conformed to the rotor blade flow path specified by NASA for compatibility with their test equipment. Fabrication was accomplished on three rotors, a bladeless rotor, a solid rotor, and an air-cooled rotor.

Streakline flow visualization study of a horseshoe vortex in a large-scale, two-dimensional turbine stator cascade

NASA Technical Reports Server (NTRS)

Gaugler, R. E.; Russell, L. M.

1979-01-01

Neutrally bouyant helium-filled bubbles were observed as they followed the streamlines in a horseshoe vortex system around the vane leading edge in a large scale, two dimensional, turbine stator cascade. Inlet Reynolds number, based on true chord, ranged between 100,000 to 300,000. Bubbles were introduced into the endwall boundary layer through a slot upstream of the vane leading edge. The paths of the bubbles were recorded photographically as streaklines on 16 mm movie film. Individual frames from the film were selected, and overlayed to show the details of the horseshoe vortex around the leading edge. The transport of the vortex across the passage near the leading edge is clearly seen when compared to the streaks formed by bubbles carried in the main stream. Limiting streamlines on the endwall surface were traced by the flow of oil drops.

Quiet Clean Short-Haul Experimental Engine (QCSEE) aerodynamic characteristics of 30.5 centimeter diameter inlets

NASA Technical Reports Server (NTRS)

Paul, D. L.

1975-01-01

A low speed test program was conducted in a 9- by 15-foot V/STOL wind tunnel to investigate internal performance characteristics and determine key design features required for an inlet to meet the demanding operational conditions of the QCSEE application. Four models each having a design average throat Mach number of 0.79 were tested over a range of incidence angle, throat Mach number, and freestream velocity. Principal design variable was internal lip diameter ratio. Stable, efficient inlet performance was found to be feasible at and beyond the 50 deg incidence angle required by the QCSEE application at its 41.2 m/sec (80 knot) nominal takeoff velocity, through suitably designed inlet lip and diffuser components. Forebody design was found to significantly impact flow stability via nose curvature. Measured inlet wall pressures were used to select a location for the inlet throat Mach number control's static pressure port that properly balanced the conflicting demands of relative insensitivity to flow incidence and sufficiently high response to changes in engine flow demand.

Investigation on centrifugal impeller in an axial-radial combined compressor with inlet distortion

NASA Astrophysics Data System (ADS)

Li, Du; Yang, Ce; Zhao, Ben; Zhou, Mi; Qi, Mingxu; Zhang, Jizhong

2011-12-01

Assembling an axial rotor and a stator at centrifugal compressor upstream to build an axial-radial combined compressor could achieve high pressure ratio and efficiency by appropriate size augment. Then upstream potential flow and wake effect appear at centrifugal impeller inlet. In this paper, the axial-radial compressor is unsteadily simulated by three-dimensional Reynolds averaged Navier-Stokes equations with uniform and circumferential distorted total pressure inlet condition to investigate upstream effect on radial rotor. The results show that span-wise nonuniform total pressure distribution is generated and radial and circumferential combined distortion is formed at centrifugal rotor inlet. The upstream stator wake deflects to rotor rotation direction and decreases with blade span increases. Circumferential distortion causes different separated flow formations at different pitch positions. The tip leakage vortex is suppressed in centrifugal blade passages. Under distorted inlet condition, flow direction of centrifugal impeller leading edge upstream varies evidently near hub and shroud but varies slightly at mid-span. In addition, compressor stage inlet distortion produces remarkable effect on blade loading of centrifugal blade both along chordwise and pitchwise.

Review and evaluation of recent developments in melic inlet dynamic flow distortion prediction and computer program documentation and user's manual estimating maximum instantaneous inlet flow distortion from steady-state total pressure measurements with full, limited, or no dynamic data

NASA Technical Reports Server (NTRS)

Schweikhard, W. G.; Dennon, S. R.

1986-01-01

A review of the Melick method of inlet flow dynamic distortion prediction by statistical means is provided. These developments include the general Melick approach with full dynamic measurements, a limited dynamic measurement approach, and a turbulence modelling approach which requires no dynamic rms pressure fluctuation measurements. These modifications are evaluated by comparing predicted and measured peak instantaneous distortion levels from provisional inlet data sets. A nonlinear mean-line following vortex model is proposed and evaluated as a potential criterion for improving the peak instantaneous distortion map generated from the conventional linear vortex of the Melick method. The model is simplified to a series of linear vortex segments which lay along the mean line. Maps generated with this new approach are compared with conventionally generated maps, as well as measured peak instantaneous maps. Inlet data sets include subsonic, transonic, and supersonic inlets under various flight conditions.

Flow Straightener for a Rotating-Drum Liquid Separator

NASA Technical Reports Server (NTRS)

O'Coin, James R.; Converse, David G.; Rethke, Donald W.

2004-01-01

A flow straightener has been incorporated into a rotary liquid separator that originally comprised an inlet tube, a shroud plate, an impeller, an inner drum, an outer drum, a housing, a pitot tube, and a hollow shaft motor. As a consequence of the original geometry of the impeller, shroud, inner drum, and hollow shaft, swirl was created in the airflow inside the hollow shaft during operation. The swirl speed was large enough to cause a significant pressure drop. The flow straightener consists of vanes on the back side of the shroud plate. These vanes compartmentalize the inside of the inner drum in such a way as to break up the flow path and thereby stop the air from swirling; as a result, the air enters the hollow shaft with a predominantly axial velocity instead of a swirl. Tests of the rotary liquid separator at an airflow rate of 10 cu ft/min (0.0047 cu m/s) revealed that the dynamic pressure drop was 8 in. of water (approx.=2 kPa) in the absence of the flow straightener and was reduced to 1 in. of water (approx.=0.25 kPa) in the presence of the flow straightener.

Study of compressible flow through a rectangular-to-semiannular transition duct

NASA Technical Reports Server (NTRS)

Foster, Jeffry; Okiishi, Theodore H.; Wendt, Bruce J.; Reichert, Bruce A.

1995-01-01

Detailed flow field measurements are presented for compressible flow through a diffusing rectangular-to-semiannular transition duct. Comparisons are made with published computational results for flow through the duct. Three-dimensional velocity vectors and total pressures were measured at the exit plane of the diffuser model. The inlet flow was also measured. These measurements are made using calibrated five-hole probes. Surface oil flow visualization and surface static pressure data were also taken. The study was conducted with an inlet Mach number of 0.786. The diffuser Reynolds based on the inlet centerline velocity and the exit diameter of the diffuser was 3,200,000. Comparison of the measured data with previously published computational results are made. Data demonstrating the ability of vortex generators to reduce flow separation and circumferential distortion is also presented.

Characterization of a Head-Only Aerosol Exposure System for Nonhuman Primates

DTIC Science & Technology

2010-01-01

Alicat Scientific). The flow used provided isokinetic sampling, assuming the presence of laminar flow at the filter inlet, so that the chamber...airflow measured at the chamber inlet using a Gilibrator flow meter (Sensidyne LP) was 29.79±0.88L/min (n= 10; CV=2.97%). With both the aerosol

Polyport atmospheric gas sampler

DOEpatents

Guggenheim, S. Frederic

1995-01-01

An atmospheric gas sampler with a multi-port valve which allows for multi, sequential sampling of air through a plurality of gas sampling tubes mounted in corresponding gas inlet ports. The gas sampler comprises a flow-through housing which defines a sampling chamber and includes a gas outlet port to accommodate a flow of gases through the housing. An apertured sample support plate defining the inlet ports extends across and encloses the sampling chamber and supports gas sampling tubes which depend into the sampling chamber and are secured across each of the inlet ports of the sample support plate in a flow-through relation to the flow of gases through the housing during sampling operations. A normally closed stopper means mounted on the sample support plate and operatively associated with each of the inlet ports blocks the flow of gases through the respective gas sampling tubes. A camming mechanism mounted on the sample support plate is adapted to rotate under and selectively lift open the stopper spring to accommodate a predetermined flow of gas through the respective gas sampling tubes when air is drawn from the housing through the outlet port.

CFD application to supersonic/hypersonic inlet airframe integration. [computational fluid dynamics (CFD)

NASA Technical Reports Server (NTRS)

Benson, Thomas J.

1988-01-01

Supersonic external compression inlets are introduced, and the computational fluid dynamics (CFD) codes and tests needed to study flow associated with these inlets are outlined. Normal shock wave turbulent boundary layer interaction is discussed. Boundary layer control is considered. Glancing sidewall shock interaction is treated. The CFD validation of hypersonic inlet configurations is explained. Scramjet inlet modules are shown.

Initial Flow Matching Results of MHD Energy Bypass on a Supersonic Turbojet Engine Using the Numerical Propulsion System Simulation (NPSS) Environment

NASA Technical Reports Server (NTRS)

Benyo, Theresa L.

2010-01-01

Preliminary flow matching has been demonstrated for a MHD energy bypass system on a supersonic turbojet engine. The Numerical Propulsion System Simulation (NPSS) environment was used to perform a thermodynamic cycle analysis to properly match the flows from an inlet to a MHD generator and from the exit of a supersonic turbojet to a MHD accelerator. Working with various operating conditions such as the enthalpy extraction ratio and isentropic efficiency of the MHD generator and MHD accelerator, interfacing studies were conducted between the pre-ionizers, the MHD generator, the turbojet engine, and the MHD accelerator. This paper briefly describes the NPSS environment used in this analysis and describes the NPSS analysis of a supersonic turbojet engine with a MHD generator/accelerator energy bypass system. Results from this study have shown that using MHD energy bypass in the flow path of a supersonic turbojet engine increases the useful Mach number operating range from 0 to 3.0 Mach (not using MHD) to an explored and desired range of 0 to 7.0 Mach.

Compressor Study to Meet Large Civil Tilt Rotor Engine Requirements

NASA Technical Reports Server (NTRS)

Veres, Joseph P.

2009-01-01

A vehicle concept study has been made to meet the requirements of the Large Civil Tilt Rotorcraft vehicle mission. A vehicle concept was determined, and a notional turboshaft engine system study was conducted. The engine study defined requirements for the major engine components, including the compressor. The compressor design-point goal was to deliver a pressure ratio of 31:1 at an inlet weight flow of 28.4 lbm/sec. To perform a conceptual design of two potential compressor configurations to meet the design requirement, a mean-line compressor flow analysis and design code were used. The first configuration is an eight-stage axial compressor. Some challenges of the all-axial compressor are the small blade spans of the rear-block stages being 0.28 in., resulting in the last-stage blade tip clearance-to-span ratio of 2.4%. The second configuration is a seven-stage axial compressor, with a centrifugal stage having a 0.28-in. impeller-exit blade span. The compressors conceptual designs helped estimate the flow path dimensions, rotor leading and trailing edge blade angles, flow conditions, and velocity triangles for each stage.

Compressor Study to Meet Large Civil Tilt Rotor Engine Requirements

NASA Technical Reports Server (NTRS)

Veres, Joseph P.

2009-01-01

A vehicle concept study has been made to meet the requirements of the Large Civil Tilt Rotorcraft vehicle mission. A vehicle concept was determined, and a notional turboshaft engine system study was conducted. The engine study defined requirements for the major engine components, including the compressor. The compressor design-point goal was to deliver a pressure ratio of 31:1 at an inlet weight flow of 28.4 lbm/sec. To perform a conceptual design of two potential compressor configurations to meet the design requirement, a mean-line compressor flow analysis and design code were used. The first configuration is an eight-stage axial compressor. Some challenges of the all-axial compressor are the small blade spans of the rear-block stages being 0.28 in., resulting in the last-stage blade tip clearance-to-span ratio of 2.4 percent. The second configuration is a seven-stage axial compressor, with a centrifugal stage having a 0.28-in. impeller-exit blade span. The compressors conceptual designs helped estimate the flow path dimensions, rotor leading and trailing edge blade angles, flow conditions, and velocity triangles for each stage.

Testing and Performance Verification of a High Bypass Ratio Turbofan Rotor in an Internal Flow Component Test Facility

NASA Technical Reports Server (NTRS)

VanZante, Dale E.; Podboy, Gary G.; Miller, Christopher J.; Thorp, Scott A.

2009-01-01

A 1/5 scale model rotor representative of a current technology, high bypass ratio, turbofan engine was installed and tested in the W8 single-stage, high-speed, compressor test facility at NASA Glenn Research Center (GRC). The same fan rotor was tested previously in the GRC 9x15 Low Speed Wind Tunnel as a fan module consisting of the rotor and outlet guide vanes mounted in a flight-like nacelle. The W8 test verified that the aerodynamic performance and detailed flow field of the rotor as installed in W8 were representative of the wind tunnel fan module installation. Modifications to W8 were necessary to ensure that this internal flow facility would have a flow field at the test package that is representative of flow conditions in the wind tunnel installation. Inlet flow conditioning was designed and installed in W8 to lower the fan face turbulence intensity to less than 1.0 percent in order to better match the wind tunnel operating environment. Also, inlet bleed was added to thin the casing boundary layer to be more representative of a flight nacelle boundary layer. On the 100 percent speed operating line the fan pressure rise and mass flow rate agreed with the wind tunnel data to within 1 percent. Detailed hot film surveys of the inlet flow, inlet boundary layer and fan exit flow were compared to results from the wind tunnel. The effect of inlet casing boundary layer thickness on fan performance was quantified. Challenges and lessons learned from testing this high flow, low static pressure rise fan in an internal flow facility are discussed.

Development and Characterization a Single-Active-Chamber Piezoelectric Membrane Pump with Multiple Passive Check Valves.

PubMed

Zhang, Ronghui; You, Feng; Lv, Zhihan; He, Zhaocheng; Wang, Haiwei; Huang, Ling

2016-12-12

In order to prevent the backward flow of piezoelectric pumps, this paper presents a single-active-chamber piezoelectric membrane pump with multiple passive check valves. Under the condition of a fixed total number of passive check valves, by means of changing the inlet valves and outlet valves' configuration, the pumping characteristics in terms of flow rate and backpressure are experimentally investigated. Like the maximum flow rate and backpressure, the testing results show that the optimal frequencies are significantly affected by changes in the number inlet valves and outlet valves. The variation ratios of the maximum flow rate and the maximum backpressure are up to 66% and less than 20%, respectively. Furthermore, the piezoelectric pump generally demonstrates very similar flow rate and backpressure characteristics when the number of inlet valves in one kind of configuration is the same as that of outlet valves in another configuration. The comparison indicates that the backflow from the pumping chamber to inlet is basically the same as the backflow from the outlet to the pumping chamber. No matter whether the number of inlet valves or the number of outlet valves is increased, the backflow can be effectively reduced. In addition, the backpressure fluctuation can be significantly suppressed with an increase of either inlet valves or outlet valves. It also means that the pump can prevent the backflow more effectively at the cost of power consumption. The pump is very suitable for conditions where more accurate flow rates are needed and wear and fatigue of check valves often occur.

Viscous analyses for flow through subsonic and supersonic intakes

NASA Technical Reports Server (NTRS)

Povinelli, Louis A.; Towne, Charles E.

1986-01-01

A parabolized Navier-Stokes code was used to analyze a number of diffusers typical of a modern inlet design. The effect of curvature of the diffuser centerline and transitioning cross sections was evaluated to determine the primary cause of the flow distortion in the duct. Results are presented for S-shaped intakes with circular and transitioning cross sections. Special emphasis is placed on verification of the analysis to accurately predict distorted flow fields resulting from pressure-driven secondary flows. The effect of vortex generators on reducing the distortion of intakes is presented. Comparisons of the experimental and analytical total pressure contours at the exit of the intake exhibit good agreement. In the case of supersonic inlets, computations of the inlet flow field reveal that large secondary flow regions may be generated just inside of the intake. These strong flows may lead to separated flow regions and cause pronounced distortions upstream of the compressor.

Effects of nose bluntness and shock-shock interactions on blunt bodies in viscous hypersonic flows

NASA Technical Reports Server (NTRS)

Singh, D. J.; Tiwari, S. N.

1990-01-01

A numerical study was conducted to investigate the effects of blunt leading edges on the viscous flow field around a hypersonic vehicle such as the proposed National Aero-Space Plane. Attention is focused on two specific regions of the flow field. In the first region, effects of nose bluntness on the forebody flow field are investigated. The second region of the flow considered is around the leading edges of the scramjet inlet. In this region, the interaction of the forebody shock with the shock produced by the blunt leading edges of the inlet compression surfaces is analyzed. Analysis of these flow regions is required to accurately predict the overall flow field as well as to get necessary information on localized zones of high pressure and intense heating. The results for the forebody flow field are discussed first, followed by the results for the shock interaction in the inlet leading edge region.

Inlet Flow Valve Engine Analyses

NASA Technical Reports Server (NTRS)

Champagne, G. A.

2004-01-01

Pratt&Whitney, under Task Order 13 of the NASA Large Engine Technology (LET) Contract, conducted a study to determine the operating characteristics, performance and weights of Inlet Flow Valve (IFV) propulsion concepts for a Mach 2.4 High Speed Civil Transport (HSCT).

Calculation of compressible flow about three-dimensional inlets with auxiliary inlets, slats and vanes by means of a panel method

NASA Technical Reports Server (NTRS)

Hess, J. L.; Friedman, D. M.; Clark, R. W.

1985-01-01

An efficient and user oriented method was constructed for calculating flow in and about complex inlet configurations. Efficiency is attained by: (1) the use of a panel method; (2) a technique of superposition for obtaining solutions at any inlet operating condition; and (3) employment of an advanced matrix iteration technique for solving large full systems of equations, including the nonlinear equations for the Kutta condition. User concerns are addressed by the provision of several novel graphical output options that yield a more complete comprehension of the flowfield than was possible previously.

Forebody and Inlet Design for the HIFiRE 2 Flight Test

NASA Technical Reports Server (NTRS)

Ferlemann, Paul G.

2008-01-01

A forebody and inlet have been designed for the HIFiRE 2 scramjet flight test. The test will explore the operating, performance, and stability characteristics of a simple hydrocarbon-fueled scramjet combustor as it transitions from dual-mode to scramjet-mode operation and during supersonic combustion at Mach 8+ flight conditions. Requirements for the compression system were derived from inlet starting and combustor inflow requirements as well as physical size constraints. The design process is described. A planar, fixed geometry, mixed compression concept was used to produce laterally uniform flow at the inlet entrance and a conservative amount of internal contraction with respect to inlet starting. A grid sensitivity study was performed so that important flow physics caused by three-dimensional shock boundary layer interactions could be captured with confidence. Results from low Mach number operability studies, nominal trajectory cases, and high dynamic pressure heat load cases are discussed. The forebody and inlet solutions provide information for on-going combustor calculations, mass capture across the trajectory for fuel system design, and surface heating rates for thermal/structural analysis. The design has a one freestream Mach number margin for inlet starting, exceeds the high Mach number combustor entrance pressure requirement, produces high quality flow at the inlet exit for all Mach numbers and vehicle attitudes in the design space, and fits inside the booster shroud.

Numerical and experimental investigation of VG flow control for a low-boom inlet

NASA Astrophysics Data System (ADS)

Rybalko, Michael

The application of vortex generators (VGs) for shock/boundary layer interaction flow control in a novel external compression, axisymmetric, low-boom concept inlet was studied using numerical and experimental methods. The low-boom inlet design features a zero-angle cowl and relaxed isentropic compression centerbody spike, resulting in defocused oblique shocks and a weak terminating normal shock. This allows reduced external gas dynamic waves at high mass flow rates but suffers from flow separation near the throat and a large hub-side boundary layer at the Aerodynamic Interface Plane (AIP), which marks the inflow to the jet engine turbo-machinery. Supersonic VGs were investigated to reduce the shock-induced flow separation near the throat while subsonic VGs were investigated to reduce boundary layer radial distortion at the AIP. To guide large-scale inlet experiments, Reynolds-Averaged Navier-Stokes (RANS) simulations using three-dimensional, structured, chimera (overset) grids and the WIND-US code were conducted. Flow control cases included conventional and novel types of vortex generators at positions both upstream of the terminating normal shock (supersonic VGs) and downstream (subsonic VGs). The performance parameters included incompressible axisymmetric shape factor, post-shock separation area, inlet pressure recovery, and mass flow ratio. The design of experiments (DOE) methodology was used to select device size and location, analyze the resulting data, and determine the optimal choice of device geometry. Based on the above studies, a test matrix of supersonic and subsonic VGs was adapted for a large-scale inlet test to be conducted at the 8'x6' supersonic wind tunnel at NASA Glenn Research Center (GRC). Comparisons of RANS simulations with data from the Fall 2010 8'x6' inlet test showed that predicted VG performance trends and case rankings for both supersonic and subsonic devices were consistent with experimental results. For example, experimental surface oil flow visualization revealed a significant post-shock separation bubble with flow recirculation for the baseline (no VG) case that was substantially broken up in the micro-ramp VG case, consistent with simulations. Furthermore, the predicted subsonic VG performance with respect to a reduction in radial distortion (quantified in terms of axisymmetric incompressible shape factor) was found to be consistent with boundary layer rake measurements. To investigate the unsteady turbulent flow features associated with the shock-induced flow separation and the hub-side boundary layer, a detached eddy simulation (DES) approach using the WIND-US code was employed to model the baseline inlet flow field. This approach yielded improved agreement with experimental data for time-averaged diffuser stagnation pressure profiles and allowed insight into the pressure fluctuations and turbulent kinetic energy distributions which may be present at the AIP. In addition, streamwise shock position statistics were obtained and compared with experimental Schlieren results. The predicted shock oscillations were much weaker than those seen experimentally (by a factor of four), which indicates that the mechanism for the experimental shock oscillations was not captured. In addition, the novel supersonic vortex generator geometries were investigated experimentally (prior to the large-scale inlet 8'x6' wind tunnel tests) in an inlet-relevant flow field containing a Mach 1.4 normal shock wave followed by a subsonic diffuser. A parametric study of device height and distance upstream of the normal shock was undertaken for split-ramp and ramped-vane geometries. Flow field diagnostics included high-speed Schlieren, oil flow visualization, and Pitot-static pressure measurements. Parameters including flow separation, pressure recovery, centerline incompressible boundary layer shape factor, and shock stability were analyzed and compared to the baseline uncontrolled case. While all vortex generators tested eliminated centerline flow separation, the presence of VGs also increased the significant three-dimensionality of the flow via increased side-wall interaction. The stronger streamwise vorticity generated by ramped-vanes also yielded improved pressure recovery and fuller boundary layer velocity profiles within the subsonic diffuser. (Abstract shortened by UMI.)

Effect of a rotating propeller on the separation angle of attack and distortion in ducted propeller inlets

NASA Technical Reports Server (NTRS)

Boldman, D. R.; Iek, C.; Hwang, D. P.; Larkin, M.; Schweiger, P.

1993-01-01

The present study represents an extension of an earlier wind tunnel experiment performed with the P&W 17-in. Advanced Ducted Propeller (ADP) Simulator operating at Mach 0.2. In order to study the effects of a rotating propeller on the inlet flow, data were obtained in the UTRC 10- by 15-Foot Large Subsonic Wind Tunnel with the same hardware and instrumentation, but with the propeller removed. These new tests were performed over a range of flow rates which duplicated flow rates in the powered simulator program. The flow through the inlet was provided by a remotely located vacuum source. A comparison of the results of this flow-through study with the previous data from the powered simulator indicated that in the conventional inlet the propeller produced an increase in the separation angle of attack between 4.0 deg at a specific flow of 22.4 lb/sec-sq ft to 2.7 deg at a higher specific flow of 33.8 lb/sec-sq ft. A similar effect on separation angle of attack was obtained by using stationary blockage rather than a propeller.

Performance of a high-work low aspect ration turbine tested with a realistic inlet radial temperature profile

NASA Technical Reports Server (NTRS)

Stabe, R. G.; Whitney, W. J.; Moffitt, T. P.

1984-01-01

Experimental results are presented for a 0.767 scale model of the first stage of a two-stage turbine designed for a high by-pass ratio engine. The turbine was tested with both uniform inlet conditions and with an inlet radial temperature profile simulating engine conditions. The inlet temperature profile was essentially mixed-out in the rotor. There was also substantial underturning of the exit flow at the mean diameter. Both of these effects were attributed to strong secondary flows in the rotor blading. There were no significant differences in the stage performance with either inlet condition when differences in tip clearance were considered. Performance was very close to design intent in both cases.

Top-mounted inlet system feasibility for transonic-supersonic fighter aircraft. [V/STOL aircraft

NASA Technical Reports Server (NTRS)

Williams, T. L.; Hunt, B. L.; Smeltzer, D. B.; Nelms, W. P.

1981-01-01

The more salient findings are presented of recent top inlet performance evaluations aimed at assessing the feasibility of top-mounted inlet systems for transonic-supersonic fighter aircraft applications. Top inlet flow field and engine-inlet performance test data show the influence of key aircraft configuration variables-inlet longitudinal position, wing leading-edge extension planform area, canopy-dorsal integration, and variable incidence canards-on top inlet performance over the Mach range of 0.6 to 2.0. Top inlet performance data are compared with those or more conventional inlet/airframe integrations in an effort to assess the viability of top-mounted inlet systems relative to conventional inlet installations.

Flow Control in a Compact Inlet

NASA Astrophysics Data System (ADS)

Vaccaro, John C.

2011-12-01

An experimental investigation of flow control, via various control jets actuators, was undertaken to eliminate separation and secondary flows in a compact inlet. The compact inlet studied was highly aggressive with a length-to-diameter ratio of 1.5. A brand new facility was designed and built to enable various actuation methodologies as well as multiple measurement techniques. Techniques included static surface pressure, total pressure, and stereoscopic particle image velocimetry. Experimental data were supplemented with numerical simulations courtesy of Prof. Kenneth Jansen, Dr. Onkar Sahni, and Yi Chen. The baseline flow field was found to be dominated by two massive separations and secondary flow structures. These secondary structures were present at the aerodynamic interface plane in the form of two counter-rotating vortices inducing upwash along centerline. A dominant shedding frequency of 350 Hz was measured both at the aerodynamic interface plane and along the lower surface of the inlet. Flow control experiments started utilizing a pair of control jets placed in streamwise locations where flow was found to separate. Tests were performed for a range of inlet Mach numbers from 0.2 to 0.44. Steady and unsteady static pressure measurements along the upper and lower walls of the duct were performed for various combinations of actuation. The parameters that were tested include the control jets momentum coefficient, their blowing ratio, the actuation frequency, as well as different combinations of jets. It was shown that using mass flux ratio as a criterion to define flow control is not sufficient, and one needs to provide both the momentum coefficient and the blowing ratio to quantify the flow control performance. A detailed study was undertaken on controlling the upstream separation point for an inlet Mach number of 0.44. Similar to the baseline flow field, the flow field associated with the activation of a two-dimensional control jet actuator was dominated by secondary flow structures. Unlike the baseline, these secondary flow structures produced downwash along the centerline. The formation of such structures was caused by the core flow stagnating on the lower surface near the aerodynamic interface plane. Using the two-dimensional steady jet resulted in an increase in the spanwise flow within the inlet and a reduction in the energy content of the 350 Hz shedding frequency. Unsteady forcing did not show much improvement over steady forcing for this configuration. A spanwise varying control jet and a hybrid Coanda jet / vortex generator jets were tested to reduce the three-dimensionality of the flow field. It was found that anytime the flow control method suppressed separation along the centerline, counter-rotating vortices existed in the lower corners of the aerodynamic interface plane.

An investigation of several NACA 1-series nose inlets with and without protruding central bodies at high-subsonic Mach numbers and at a Mach number of 1.2

NASA Technical Reports Server (NTRS)

Pendley, Robert E; Robinson, Harold L

1950-01-01

An investigation of three NACA 1-series nose inlets, two of which were fitted with protruded central bodies, was conducted in the Langley 8-foot high-speed tunnel. An elliptical-nose body, which had a critical Mach number approximately equal to that of one of the nose inlets, was also tested. Tests were made near zero angle of attack for a Mach number range from 0.4 to 0.925 and for the supersonic Mach number of 1.2. The inlet-velocity-ratio range extended from zero to a maximum value of 1.34. Measurements included pressure distribution, external drag, and total-pressure loss of the internal flow near the inlet. Drag was not measured for the tests at the supersonic Mach number. Over the range of inlet-velocity ratio investigated, the calculated external pressure-drag coefficient at a Mach number of 1.2 was consecutively lower for the nose inlets of higher critical Mach number, and the pressure-drag coefficient of the longest nose inlet was in the range of pressure-drag coefficient for two solid noses of fineness ratio 2.4 and 6.0. For Mach numbers below the Mach number of the supercritical drag rise, extrapolation of the test data indicated that the external drag of the nose inlets was little affected by the addition of central bodies at or slightly below the minimum inlet-velocity ratio for unseparated central-body flow. The addition of central bodies to the nose inlets also led to no appreciable effects on either the Mach number of the supercritical drag rise, or, for inlet-velocity ratios high enough to avoid a pressure peak at the inlet lip, on the critical Mach number. The total-pressure recovery of the inlets tested, which were of a subsonic type, was sensibly unimpaired at the supersonic Mach number of 1.2 Low-speed measurements of the minimum inlet-velocity ratio for unseparated central-body flow appear to be applicable for Mach numbers extending to 1.2.

Control of the Development of Swirling Airflow Dynamics and Its Impact on Biomass Combustion Characteristics

NASA Astrophysics Data System (ADS)

Barmina, I.; Valdmanis, R.; Zaķe, M.

2017-06-01

The development of the swirling flame flow field and gasification/ combustion dynamics at thermo-chemical conversion of biomass pellets has experimentally been studied using a pilot device, which combines a biomass gasifier and combustor by varying the inlet conditions of the fuel-air mixture into the combustor. Experimental modelling of the formation of the cold nonreacting swirling airflow field above the inlet nozzle of the combustor and the upstream flow formation below the inlet nozzle has been carried out to assess the influence of the inlet nozzle diameter, as well primary and secondary air supply rates on the upstream flow formation and air swirl intensity, which is highly responsible for the formation of fuel-air mixture entering the combustor and the development of combustion dynamics downstream of the combustor. The research results demonstrate that at equal primary axial and secondary swirling air supply into the device a decrease in the inlet nozzle diameter enhances the upstream air swirl formation by increasing swirl intensity below the inlet nozzle of the combustor. This leads to the enhanced mixing of the combustible volatiles with the air swirl below the inlet nozzle of the combustor providing a more complete combustion of volatiles and an increase in the heat output of the device.

A One Dimensional, Time Dependent Inlet/Engine Numerical Simulation for Aircraft Propulsion Systems

NASA Technical Reports Server (NTRS)

Garrard, Doug; Davis, Milt, Jr.; Cole, Gary

1999-01-01

The NASA Lewis Research Center (LeRC) and the Arnold Engineering Development Center (AEDC) have developed a closely coupled computer simulation system that provides a one dimensional, high frequency inlet/engine numerical simulation for aircraft propulsion systems. The simulation system, operating under the LeRC-developed Application Portable Parallel Library (APPL), closely coupled a supersonic inlet with a gas turbine engine. The supersonic inlet was modeled using the Large Perturbation Inlet (LAPIN) computer code, and the gas turbine engine was modeled using the Aerodynamic Turbine Engine Code (ATEC). Both LAPIN and ATEC provide a one dimensional, compressible, time dependent flow solution by solving the one dimensional Euler equations for the conservation of mass, momentum, and energy. Source terms are used to model features such as bleed flows, turbomachinery component characteristics, and inlet subsonic spillage while unstarted. High frequency events, such as compressor surge and inlet unstart, can be simulated with a high degree of fidelity. The simulation system was exercised using a supersonic inlet with sixty percent of the supersonic area contraction occurring internally, and a GE J85-13 turbojet engine.

Understanding processes controlling sediment transports at the mouth of a highly energetic inlet system (San Francisco Bay, CA)

USGS Publications Warehouse

Elias, Edwin P.L.; Hansen, Jeff E.; Barnard, P.L.; Jaffee, B.E.; Schoellhamer, D.H.

2013-01-01

San Francisco Bay is one of the largest estuaries along the U.S. West Coast and is linked to the Pacific Ocean through the Golden Gate, a 100 m deep bedrock inlet. A coupled wave, flow and sediment transport model is used to quantify the sediment linkages between San Francisco Bay, the Golden Gate, and the adjacent open coast. Flow and sediment transport processes are investigated using an ensemble average of 24 climatologically derived wave cases and a 24.8 h representative tidal cycle. The model simulations show that within the inlet, flow and sediment transport is tidally dominated and driven by asymmetry of the ebb and flood tides. Peak ebb velocities exceed the peak flood velocities in the narrow Golden Gate channel as a result of flow convergence and acceleration. Persistent flow and sediment gyres at the headland tips are formed that limit sediment transfer from the ebb-tidal delta to the inlet and into the bay. The residual transport pattern in the inlet is dominated by a lateral segregation with a large ebb-dominant sediment transport (and flow) prevailing along the deeper north side of the Golden Gate channel, and smaller flood dominant transports along the shallow southern margin. The seaward edge of the ebb-tidal delta largely corresponds to the seaward extent of strong tidal flows. On the ebb-tidal delta, both waves and tidal forcing govern flow and sediment transport. Wave focusing by the ebb-tidal delta leads to strong patterns of sediment convergence and divergence along the adjacent Ocean Beach.

Optimal Design of Passive Flow Control for a Boundary-Layer-Ingesting Offset Inlet Using Design-of-Experiments

NASA Technical Reports Server (NTRS)

Allan, Brian G.; Owens, Lewis R.; Lin, John C.

2006-01-01

This research will investigate the use of Design-of-Experiments (DOE) in the development of an optimal passive flow control vane design for a boundary-layer-ingesting (BLI) offset inlet in transonic flow. This inlet flow control is designed to minimize the engine fan-face distortion levels and first five Fourier harmonic half amplitudes while maximizing the inlet pressure recovery. Numerical simulations of the BLI inlet are computed using the Reynolds-averaged Navier-Stokes (RANS) flow solver, OVERFLOW, developed at NASA. These simulations are used to generate the numerical experiments for the DOE response surface model. In this investigation, two DOE optimizations were performed using a D-Optimal Response Surface model. The first DOE optimization was performed using four design factors which were vane height and angles-of-attack for two groups of vanes. One group of vanes was placed at the bottom of the inlet and a second group symmetrically on the sides. The DOE design was performed for a BLI inlet with a free-stream Mach number of 0.85 and a Reynolds number of 2 million, based on the length of the fan-face diameter, matching an experimental wind tunnel BLI inlet test. The first DOE optimization required a fifth order model having 173 numerical simulation experiments and was able to reduce the DC60 baseline distortion from 64% down to 4.4%, while holding the pressure recovery constant. A second DOE optimization was performed holding the vanes heights at a constant value from the first DOE optimization with the two vane angles-of-attack as design factors. This DOE only required a second order model fit with 15 numerical simulation experiments and reduced DC60 to 3.5% with small decreases in the fourth and fifth harmonic amplitudes. The second optimal vane design was tested at the NASA Langley 0.3- Meter Transonic Cryogenic Tunnel in a BLI inlet experiment. The experimental results showed a 80% reduction of DPCP(sub avg), the circumferential distortion level at the engine fan-face.

Optimal Design of Passive Flow Control for a Boundary-Layer-Ingesting Offset Inlet Using Design-of-Experiments

NASA Technical Reports Server (NTRS)

Allan, Brian G.; Owens, Lewis R., Jr.; Lin, John C.

2006-01-01

This research will investigate the use of Design-of-Experiments (DOE) in the development of an optimal passive flow control vane design for a boundary-layer-ingesting (BLI) offset inlet in transonic flow. This inlet flow control is designed to minimize the engine fan face distortion levels and first five Fourier harmonic half amplitudes while maximizing the inlet pressure recovery. Numerical simulations of the BLI inlet are computed using the Reynolds-averaged Navier-Stokes (RANS) flow solver, OVERFLOW, developed at NASA. These simulations are used to generate the numerical experiments for the DOE response surface model. In this investigation, two DOE optimizations were performed using a D-Optimal Response Surface model. The first DOE optimization was performed using four design factors which were vane height and angles-of-attack for two groups of vanes. One group of vanes was placed at the bottom of the inlet and a second group symmetrically on the sides. The DOE design was performed for a BLI inlet with a free-stream Mach number of 0.85 and a Reynolds number of 2 million, based on the length of the fan face diameter, matching an experimental wind tunnel BLI inlet test. The first DOE optimization required a fifth order model having 173 numerical simulation experiments and was able to reduce the DC60 baseline distortion from 64% down to 4.4%, while holding the pressure recovery constant. A second DOE optimization was performed holding the vanes heights at a constant value from the first DOE optimization with the two vane angles-of-attack as design factors. This DOE only required a second order model fit with 15 numerical simulation experiments and reduced DC60 to 3.5% with small decreases in the fourth and fifth harmonic amplitudes. The second optimal vane design was tested at the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel in a BLI inlet experiment. The experimental results showed a 80% reduction of DPCPavg, the circumferential distortion level at the engine fan face.

Methodology for the Design of Streamline-Traced External-Compression Supersonic Inlets

NASA Technical Reports Server (NTRS)

Slater, John W.

2014-01-01

A design methodology based on streamline-tracing is discussed for the design of external-compression, supersonic inlets for flight below Mach 2.0. The methodology establishes a supersonic compression surface and capture cross-section by tracing streamlines through an axisymmetric Busemann flowfield. The compression system of shock and Mach waves is altered through modifications to the leading edge and shoulder of the compression surface. An external terminal shock is established to create subsonic flow which is diffused in the subsonic diffuser. The design methodology was implemented into the SUPIN inlet design tool. SUPIN uses specified design factors to design the inlets and computes the inlet performance, which includes the flow rates, total pressure recovery, and wave drag. A design study was conducted using SUPIN and the Wind-US computational fluid dynamics code to design and analyze the properties of two streamline-traced, external-compression (STEX) supersonic inlets for Mach 1.6 freestream conditions. The STEX inlets were compared to axisymmetric pitot, two-dimensional, and axisymmetric spike inlets. The STEX inlets had slightly lower total pressure recovery and higher levels of total pressure distortion than the axisymmetric spike inlet. The cowl wave drag coefficients of the STEX inlets were 20% of those for the axisymmetric spike inlet. The STEX inlets had external sound pressures that were 37% of those of the axisymmetric spike inlet, which may result in lower adverse sonic boom characteristics. The flexibility of the shape of the capture cross-section may result in benefits for the integration of STEX inlets with aircraft.

High-speed pulsed mixing in a short distance with high-frequency switching of pumping from three inlets

NASA Astrophysics Data System (ADS)

Sugano, K.; Nakata, A.; Tsuchiya, T.; Tabata, O.

2015-08-01

In this study, we propose a mixing method using alternate pulsed flows from three inlets with flow direction control. In conventional pulsed mixing, a residual flow near the sidewalls inhibits the rapid mixing of two solutions at high switching frequency. In this study, we addressed this issue in order to perform rapid mixing in a short distance with a low Reynolds number. We fabricated a microfluidic mixing device consisting of a cross-shaped mixing channel with three inlet microchannels and three valveless micropumps. In conventional T-shaped or Y-shaped mixing channels, a residual flow is observed because of the incomplete switching of solutions. The three inlet configuration enabled us to split the residual flow at a switching frequency of pumping of up to 200 Hz, thus resulting in rapid mixing. Furthermore, by controlling the flow direction at the confluent area using the reverse flow of the micropump, the mixing speed was dramatically increased because of the complete switching of the two solutions. As a result, we achieved the mixing time of 3.6 ms and the mixing length of 20.7 µm, which were necessary to achieve a 90% mixing ratio at a high micropump switching frequency of 400 Hz and reverse flow ratio of 1/4.

The effects of micro-vortex generators on normal shock wave/boundary layer interactions

NASA Astrophysics Data System (ADS)

Herges, Thomas G.

Shock wave/boundary-layer interactions (SWBLIs) are complex flow phenomena that are important in the design and performance of internal supersonic and transonic flow fields such as engine inlets. This investigation was undertaken to study the effects of passive flow control devices on normal shock wave/boundary layer interactions in an effort to gain insight into the physics that govern these complex interactions. The work concentrates on analyzing the effects of vortex generators (VGs) as a flow control method by contributing a greater understanding of the flowfield generated by these devices and characterizing their effects on the SWBLI. The vortex generators are utilized with the goal of improving boundary layer health (i.e., reducing/increasing the boundary-layer incompressible shape factor/skin friction coefficient) through a SWBLI, increasing pressure recovery, and reducing flow distortion at the aerodynamic interface plane while adding minimal drag to the system. The investigation encompasses experiments in both small-scale and large-scale inlet testing, allowing multiple test beds for improving the characterization and understanding of vortex generators. Small-scale facility experiments implemented instantaneous schlieren photography, surface oil-flow visualization, pressure-sensitive paint, and particle image velocimetry to characterize the effects of an array of microramps on a normal shock wave/boundary-layer interaction. These diagnostics measured the time-averaged and instantaneous flow organization in the vicinity of the microramps and SWBLI. The results reveal that a microramp produces a complex vortex structure in its wake with two primary counter-rotating vortices surrounded by a train of Kelvin- Helmholtz (K-H) vortices. A streamwise velocity deficit is observed in the region of the primary vortices in addition to an induced upwash/downwash which persists through the normal shock with reduced strength. The microramp flow control also increased the spanwise-averaged skin-friction coefficient and reduced the spanwise-averaged incompressible shape factor, thereby improving the health of the boundary layer. The velocity in the near-wall region appears to be the best indicator of microramp effectiveness at controlling SWBLIs. Continued analysis of additional micro-vortex generator designs in the small-scale facility revealed reduced separation within a subsonic diffuser downstream of the normal shock wave/boundary layer interaction. The resulting attached flow within the diffuser from the micro-vortex generator control devices reduces shock wave position and pressure RMS fluctuations within the diffuser along with increased pressure recovery through the shock and at the entrance of the diffuser. The largest effect was observed by the micro-vortex generators that produce the strongest streamwise vortices. High-speed pressure measurements also indicated that the vortex generators shift the energy of the pressure fluctuations to higher frequencies. Implementation of micro-vortex generators into a large-scale, supersonic, axisymmetric, relaxed-compression inlet have been investigated with the use of a unique and novel flow-visualization measurement system designed and successfully used for the analysis of both upstream micro-VGs (MVGs) and downstream VGs utilizing surface oil-flow visualization and pressure-sensitive paint measurements. The inlet centerbody and downstream diffuser vortex-generator regions were imaged during wind-tunnel testing internally through the inlet cowl with the diagnostic system attached to the cowl. Surface-flow visualization revealed separated regions along the inlet centerbody for large mass-flow rates without vortex generators. Upstream vortex generators did reduce separation in the subsonic diffuser, and a unique perspective of the flowfield produced by the downstream vortex generators was obtained. In addition, pressure distributions on the inlet centerbody and vortex generators were measured with pressure-sensitive paint. At low mass-flow ratios the onset of buzz occurs in the large-scale low-boom inlet. Inlet buzz and how it is affected by vortex generators was characterized using shock tracking through high-speed schlieren imaging and pressure fluctuation measurements. The analysis revealed a dominant low frequency oscillation at 21.0 Hz for the single-stream inlet, corresponding with the duration of one buzz cycle. Pressure oscillations prior to the onset of buzz were not detected, leaving the location where the shock wave triggers large separation on the compression spike as the best indicator for the onset of buzz. The driving mechanism for a buzz cycle has been confirmed as the rate of depressurization and repressurization of the inlet as the buzz cycle fluctuates between an effectively unstarted (blocked) inlet and supercritical operation (choked flow), respectively. High-frequency shock position oscillations/pulsations (spike buzz) were also observed throughout portions of the inlet buzz cycle. The primary effect of the VGs was to trigger buzz at a higher mass-flow ratio.

IPAC-Inlet Performance Analysis Code

NASA Technical Reports Server (NTRS)

Barnhart, Paul J.

1997-01-01

A series of analyses have been developed which permit the calculation of the performance of common inlet designs. The methods presented are useful for determining the inlet weight flows, total pressure recovery, and aerodynamic drag coefficients for given inlet geometric designs. Limited geometric input data is required to use this inlet performance prediction methodology. The analyses presented here may also be used to perform inlet preliminary design studies. The calculated inlet performance parameters may be used in subsequent engine cycle analyses or installed engine performance calculations for existing uninstalled engine data.

Water Flow Testing and Unsteady Pressure Analysis of a Two-Bladed Liquid Oxidizer Pump Inducer

NASA Technical Reports Server (NTRS)

Schwarz, Jordan B.; Mulder, Andrew; Zoladz, Thomas

2011-01-01

The unsteady fluid dynamic performance of a cavitating two-bladed oxidizer turbopump inducer was characterized through sub-scale water flow testing. While testing a novel inlet duct design that included a cavitation suppression groove, unusual high-frequency pressure oscillations were observed. With potential implications for inducer blade loads, these high-frequency components were analyzed extensively in order to understand their origins and impacts to blade loading. Water flow testing provides a technique to determine pump performance without the costs and hazards associated with handling cryogenic propellants. Water has a similar density and Reynolds number to liquid oxygen. In a 70%-scale water flow test, the inducer-only pump performance was evaluated. Over a range of flow rates, the pump inlet pressure was gradually reduced, causing the flow to cavitate near the pump inducer. A nominal, smooth inducer inlet was tested, followed by an inlet duct with a circumferential groove designed to suppress cavitation. A subsequent 52%-scale water flow test in another facility evaluated the combined inducer-impeller pump performance. With the nominal inlet design, the inducer showed traditional cavitation and surge characteristics. Significant bearing loads were created by large side loads on the inducer during synchronous cavitation. The grooved inlet successfully mitigated these loads by greatly reducing synchronous cavitation, however high-frequency pressure oscillations were observed over a range of frequencies. Analytical signal processing techniques showed these oscillations to be created by a rotating, multi-celled train of pressure pulses, and subsequent CFD analysis suggested that such pulses could be created by the interaction of rotating inducer blades with fluid trapped in a cavitation suppression groove. Despite their relatively low amplitude, these high-frequency pressure oscillations posed a design concern due to their sensitivity to flow conditions and test scale. The amplitude and frequency of oscillations varied considerably over the pump s operating space, making it difficult to predict blade loads.

Two-stage, low noise advanced technology fan. 4: Aerodynamic final report

NASA Technical Reports Server (NTRS)

Harley, K. G.; Keenan, M. J.

1975-01-01

A two-stage research fan was tested to provide technology for designing a turbofan engine for an advanced, long range commercial transport having a cruise Mach number of 0.85 -0.9 and a noise level 20 EPNdB below current requirements. The fan design tip speed was 365.8m/sec (1200ft/sec);the hub/tip ratio was 0.4; the design pressure ratio was 1.9; and the design specific flow was 209.2 kg/sec/sq m(42.85lbm/sec/sq ft). Two fan-versions were tested: a baseline configuration, and an acoustically treated configuration with a sonic inlet device. The baseline version was tested with uniform inlet flow and with tip-radial and hub-radial inlet flow distortions. The baseline fan with uniform inlet flow attained an efficiency of 86.4% at design speed, but the stall margin was low. Tip-radial distortion increased stall margin 4 percentage points at design speed and reduced peak efficiency one percentage point. Hub-radial distortion decreased stall margin 4 percentage points at all speeds and reduced peak efficiency at design speed 8 percentage points. At design speed, the sonic inlet in the cruise position reduced stall margin one percentage point and efficiency 1.5 to 4.5 percentage points. The sonic inlet in the approach position reduced stall margin 2 percentage points.

Investigation on heat transfer characteristics and flow performance of Methane at supercritical pressures

NASA Astrophysics Data System (ADS)

Xian, Hong Wei; Oumer, A. N.; Basrawi, F.; Mamat, Rizalman; Abdullah, A. A.

2018-04-01

The aim of this study is to investigate the heat transfer and flow characteristic of cryogenic methane in regenerative cooling system at supercritical pressures. The thermo-physical properties of supercritical methane were obtained from the National institute of Standards and Technology (NIST) webbook. The numerical model was developed based on the assumptions of steady, turbulent and Newtonian flow. For mesh independence test and model validation, the simulation results were compared with published experimental results. The effect of four different performance parameter ranges namely inlet pressure (5 to 8 MPa), inlet temperature (120 to 150 K), heat flux (2 to 5 MW/m2) and mass flux (7000 to 15000 kg/m2s) on heat transfer and flow performances were investigated. It was found that the simulation results showed good agreement with experimental data with maximum deviation of 10 % which indicates the validity of the developed model. At low inlet temperature, the change of specific heat capacity at near-wall region along the tube length was not significant while the pressure drop registered was high. However, significant variation was observed for the case of higher inlet temperature. It was also observed that the heat transfer performance and pressure drop penalty increased when the mass flux was increased. Regarding the effect of inlet pressure, the heat transfer performance and pressure drop results decreased when the inlet pressure is increased.

Vortex Generators in a Streamline-Traced, External-Compression Supersonic Inlet

NASA Technical Reports Server (NTRS)

Baydar, Ezgihan; Lu, Frank K.; Slater, John W.; Trefny, Charles J.

2017-01-01

Vortex generators within a streamline-traced, external-compression supersonic inlet for Mach 1.66 were investigated to determine their ability to increase total pressure recovery and reduce total pressure distortion. The vortex generators studied were rectangular vanes arranged in counter-rotating and co-rotating arrays. The vane geometric factors of interest included height, length, spacing, angle-of-incidence, and positions upstream and downstream of the inlet terminal shock. The flow through the inlet was simulated numerically through the solution of the steady-state, Reynolds-averaged Navier-Stokes equations on multi-block, structured grids using the Wind-US flow solver. The vanes were simulated using a vortex generator model. The inlet performance was characterized by the inlet total pressure recovery and the radial and circumferential total pressure distortion indices at the engine face. Design of experiments and statistical analysis methods were applied to quantify the effect of the geometric factors of the vanes and search for optimal vane arrays. Co-rotating vane arrays with negative angles-of-incidence positioned on the supersonic diffuser were effective in sweeping low-momentum flow from the top toward the sides of the subsonic diffuser. This distributed the low-momentum flow more evenly about the circumference of the subsonic diffuser and reduced distortion. Co-rotating vane arrays with negative angles-of-incidence or counter-rotating vane arrays positioned downstream of the terminal shock were effective in mixing higher-momentum flow with lower-momentum flow to increase recovery and decrease distortion. A strategy of combining a co-rotating vane array on the supersonic diffuser with a counter-rotating vane array on the subsonic diffuser was effective in increasing recovery and reducing distortion.

Experimental Assessment of the Hydraulics of a Miniature Axial-Flow Left Ventricular Assist Device

NASA Astrophysics Data System (ADS)

Smith, P. Alex; Cohn, William; Metcalfe, Ralph

2017-11-01

A minimally invasive partial-support left ventricular assist device (LVAD) has been proposed with a flow path from the left atrium to the arterial system to reduce left ventricular stroke work. In LVAD design, peak and average efficiency must be balanced over the operating range to reduce blood trauma. Axial flow pumps have many geometric parameters. Until recently, testing all these parameters was impractical, but modern 3D printing technology enables multi-parameter studies. Following theoretical design, experimental hydraulic evaluation in steady state conditions examines pressure, flow, pressure-flow gradient, efficiency, torque, and axial force as output parameters. Preliminary results suggest that impeller blades and stator vanes with higher inlet angles than recommended by mean line theory (MLT) produce flatter gradients and broader efficiency curves, increasing compatibility with heart physiology. These blades also produce less axial force, which reduces bearing load. However, they require slightly higher torque, which is more demanding of the motor. MLT is a low order, empirical model developed on large pumps. It does not account for the significant viscous losses in small pumps like LVADs. This emphasizes the importance of experimental testing for hydraulic design. Roderick D MacDonald Research Fund.

Wake orientation and its influence on the performance of diffusers with inlet distortion

NASA Astrophysics Data System (ADS)

Coffman, Jesse M.

Distortion at the inlet to diffusers is very common in internal flow applications. Inlet velocity distortion influences the pressure recovery and flow regimes of diffusers. This work introduced a centerline wake at the square inlet of a plane wall diffuser in two orthogonal orientations to investigate its influence on the diffuser performance. Two different wakes were generated. One was from a mesh strip which produced a velocity deficit with low turbulence intensity and two shear layers. The other wake generator was a D-shaped cylinder which produced a wake with high turbulence intensity and large length scales. These inlet conditions were generated for a diffuser with a diffusion angle of 3° and 6°. A pair of RANS simulations were used to investigate the influence of the orthogonal inlet orientations on the solution. The inlet conditions were taken from the inlet velocity field measured for the mesh strip. The flow development and exit conditions showed some similarities and some differences with the experimental results. The performance of a diffuser is typically measured through the static pressure recovery coefficient and the total pressure losses. The definition of these metrics commonly found in the literature were insufficient to discern differences between the wake orientations. New metrics were derived using the momentum flux profile parameter which related the static pressure recovery, the total pressure losses, and the velocity uniformity at the inlet and exit of the diffuser. These metrics revealed a trade-off between the total pressure losses and the uniformity of the velocity field.

Spiral inlets for steam turbines

NASA Astrophysics Data System (ADS)

Škach, Radek; Uher, Jan

2017-09-01

This paper deals with the design process of special nozzle blades for spiral inlets. Spiral inlets are used for the first stages of high pressure and intermediate pressure steam turbines with both reaction and impulse blades when throttling or sliding pressure control is applied. They improve the steam flow uniformity from the inlet pipe and thus decrease the aerodynamic losses. The proposed evaluation of the inlet angle is based on the free vortex law.

High Reynolds Number Investigation of a Flush Mounted, S-Duct Inlet With Large Amounts of Boundary Layer Ingestion

NASA Technical Reports Server (NTRS)

Berrier, Bobby L.; Carter, Melissa B.; Allan, Brian G.

2005-01-01

An experimental investigation of a flush-mounted, S-duct inlet with large amounts of boundary layer ingestion has been conducted at Reynolds numbers up to full scale. The study was conducted in the NASA Langley Research Center 0.3-Meter Transonic Cryogenic Tunnel. In addition, a supplemental computational study on one of the inlet configurations was conducted using the Navier-Stokes flow solver, OVERFLOW. Tests were conducted at Mach numbers from 0.25 to 0.83, Reynolds numbers (based on aerodynamic interface plane diameter) from 5.1 million to 13.9 million (full-scale value), and inlet mass-flow ratios from 0.29 to 1.22, depending on Mach number. Results of the study indicated that increasing Mach number, increasing boundary layer thickness (relative to inlet height) or ingesting a boundary layer with a distorted profile decreased inlet performance. At Mach numbers above 0.4, increasing inlet airflow increased inlet pressure recovery but also increased distortion. Finally, inlet distortion was found to be relatively insensitive to Reynolds number, but pressure recovery increased slightly with increasing Reynolds number.This CD-ROM supplement contains inlet data including: Boundary layer data, Duct static pressure data, performance-AIP (fan face) data, Photos, Tunnel wall P-PTO data and definitions.

Assessing Fan Flutter Stability in Presence of Inlet Distortion Using One-Way and Two-Way Coupled Methods

NASA Technical Reports Server (NTRS)

Herrick, Gregory P.

2014-01-01

Concerns regarding noise, propulsive efficiency, and fuel burn are inspiring aircraft designs wherein the propulsive turbomachines are partially (or fully) embedded within the airframe; such designs present serious concerns with regard to aerodynamic and aeromechanic performance of the compression system in response to inlet distortion. Previously, a preliminary design of a forward-swept high-speed fan exhibited flutter concerns in clean-inlet flows, and the present author then studied this fan further in the presence of off-design distorted in-flows. Continuing this research, a three-dimensional, unsteady, Navier-Stokes computational fluid dynamics code is again applied to analyze and corroborate fan performance with clean inlet flow and now with a simplified, sinusoidal distortion of total pressure at the aerodynamic interface plane. This code, already validated in its application to assess aerodynamic damping of vibrating blades at various flow conditions using a one-way coupled energy-exchange approach, is modified to include a two-way coupled timemarching aeroelastic simulation capability. The two coupling methods are compared in their evaluation of flutter stability in the presence of distorted in-flows.

Two-stage fan. 2: Data and performance with redesigned second stage rotor uniform and distorted inlet flows

NASA Technical Reports Server (NTRS)

Messenger, H. E.; Keenan, M. J.

1974-01-01

A two-stage fan with a first rotor tip speed of 1450 ft/sec (441.96 m/sec) and no inlet guide vanes was tested with uniform and distorted inlet flows, with a redesigned second rotor having a part span shroud to prevent flutter, with variable-stagger stators set in nominal positions, and without rotor casing treatment. The fan achieved a pressure ratio 2.8 at a corrected flow of 185.4 lbm/sec (84.0 kg/sec), an adiabatic efficiency of 85.0 percent, and a stall margin of 12 percent. The redesigned second rotor did not flutter. Tip radial distortion reduced the stall margin at intermediate speed, but had little effect on stall margin at high or low speeds. Hub radial distortion reduced the stall margin at design speed but increased stall margin at low speed. Circumferential distortion reduced stall pressure ratio and flow to give approximately the same stall lines with uniform inlet flow. Distortions were attenuated by the fan. For Vol. 1, see N74-11421.

Convective heat transfer studies at high temperatures with pressure gradient for inlet flow Mach number of 0.45

NASA Technical Reports Server (NTRS)

Pedrosa, A. C. F.; Nagamatsu, H. T.; Hinckel, J. A.

1984-01-01

Heat transfer measurements were determined for a flat plate with and without pressure gradient for various free stream temperatures, wall temperature ratios, and Reynolds numbers for an inlet flow Mach number of 0.45, which is a representative inlet Mach number for gas turbine rotor blades. A shock tube generated the high temperature and pressure air flow, and a variable geometry test section was used to produce inlet flow Mach number of 0.45 and accelerate the flow over the plate to sonic velocity. Thin-film platinum heat gages recorded the local heat flux for laminar, transition, and turbulent boundary layers. The free stream temperatures varied from 611 R (339 K) to 3840 R (2133 K) for a T(w)/T(r,g) temperature ratio of 0.87 to 0.14. The Reynolds number over the heat gages varied from 3000 to 690,000. The experimental heat transfer data were correlated with laminar and turbulent boundary layer theories for the range of temperatures and Reynolds numbers and the transition phenomenon was examined.

Acoustically enhanced heat exchange and drying apparatus

DOEpatents

Bramlette, T. Tazwell; Keller, Jay O.

1989-01-01

A heat transfer apparatus includes a first chamber having a first heat transfer gas inlet, a second heat transfer gas inlet, and an outlet. A first heat transfer gas source provides a first gas flow to the first chamber through the first heat transfer gas inlet. A second gas flow through a second chamber connected to the side of the first chamber, generates acoustic waves which bring about acoustical coupling of the first and second gases in the acoustically augmented first chamber. The first chamber may also include a material inlet for receiving material to be dried, in which case the gas outlet serves as a dried material and gas outlet.

Performance of a high-work, low-aspect-ratio turbine stator tested with a realistic inlet radial temperature gradient

NASA Technical Reports Server (NTRS)

Stabe, Roy G.; Schwab, John R.

1991-01-01

A 0.767-scale model of a turbine stator designed for the core of a high-bypass-ratio aircraft engine was tested with uniform inlet conditions and with an inlet radial temperature profile simulating engine conditions. The principal measurements were radial and circumferential surveys of stator-exit total temperature, total pressure, and flow angle. The stator-exit flow field was also computed by using a three-dimensional Navier-Stokes solver. Other than temperature, there were no apparent differences in performance due to the inlet conditions. The computed results compared quite well with the experimental results.

Experimental investigation of a Mach 6 fixed-geometry inlet featuring a swept external-internal compression flow field

NASA Technical Reports Server (NTRS)

Torrence, M. G.

1975-01-01

An investigation of a fixed-geometry, swept external-internal compression inlet was conducted at a Mach number of 6.0 and a test-section Reynolds number of 1.55 x 10 to the 7th power per meter. The test conditions was constant for all runs with stagnation pressure and temperature at 20 atmospheres and 500 K, respectively. Tests were made at angles of attack of -5 deg, 0 deg, 3 deg, and 5 deg. Measurements consisted of pitot- and static-pressure surveys in inlet throat, wall static pressures, and surface temperatures. Boundary-layer bleed was provided on the centerbody and on the cowl internal surface. The inlet performance was consistently high over the range of the angle of attack tested, with an overall average total pressure recovery of 78 percent and corresponding adiabatic kinetic-energy efficiency of 99 percent. The inlet throat flow distribution was uniform and the Mach number and pressure level were of the correct magnitude for efficient combustor design. The utilization of a swept compression field to meet the starting requirements of a fixed-geometry inlet produced neither flow instability nor a tendency to unstart.

Numerical Investigation of Vortex Generator Flow Control for External-Compression Supersonic Inlets

NASA Astrophysics Data System (ADS)

Baydar, Ezgihan

Vortex generators (VGs) within external-compression supersonic inlets for Mach 1.6 were investigated to determine their ability to increase total pressure recovery and reduce total pressure distortion. Ramp and vane-type VGs were studied. The geometric factors of interest included height, length, spacing, angle-of-incidence, and positions upstream and downstream of the inlet terminal shock. The flow through the inlet was simulated numerically through the solution of the steady-state, Reynolds-averaged Navier-Stokes equations on multi-block, structured grids using the Wind-US flow solver. The inlet performance was characterized by the inlet total pressure recovery and the radial and circumferential total pressure distortion indices at the engine face. Previous research of downstream VGs in the low-boom supersonic inlet demonstrated improvement in radial distortion up to 24% while my work on external-compression supersonic inlets improved radial distortion up to 86%, which is significant. The design of experiments and statistical analysis methods were applied to quantify the effect of the geometric factors of VGs and search for optimal VG arrays. From the analysis, VG angle-of-incidence and VG height were the most influential factors in increasing total pressure recovery and reducing distortion. The study on the two-dimensional external-compression inlet determined which passive flow control devices, such as counter-rotating vanes or ramps, reduce high distortion levels and improve the health of the boundary layer, relative to the baseline. Downstream vanes demonstrate up to 21% improvement in boundary layer health and 86% improvement in radial distortion. Upstream vanes demonstrated up to 3% improvement in boundary layer health and 9% improvement in radial distortion. Ramps showed no improvement in boundary layer health and radial distortion. Micro-VGs were preferred for their reduced viscous drag and improvement in total pressure recovery at the AIP. Although traditional VGs energize the flow with stronger vortex structures compared to micro-VGs, the AIP is affected with overwhelming amounts of reduced and enhanced flow regions. In summary, vanes are exceptional in reducing radial distortion and improving the health of the boundary layer compared to the ramps. In the study of the STEX inlet, vane-type vortex generators were the preferred devices for boundary layer flow control. In the supersonic diffuser, co-rotating vane arrays and counter-rotating vane arrays did not show improvement. In the subsonic diffuser, co-rotating vane arrays with negative angles-of-incidence and counter-rotating vane arrays were exceptional in reducing radial distortion and improving total pressure recovery. Downstream co-rotating vanes demonstrated up to 41% improvement in radial distortion whereas downstream counter-rotating vanes demonstrated up to 73% improvement. For downstream counter-rotating vanes, a polynomial trend between VG height and radial distortion indicate that increasing VG height improves inlet distortion. In summary, downstream vanes are exceptional in improving total pressure recovery compared to upstream vanes.

Comparison of predicted and measured low-speed performance of two 51 centimeter-diameter inlets at incidence angle

NASA Technical Reports Server (NTRS)

Albers, J. A.

1973-01-01

Theoretical and experimental internal flow characteristics of two 51-cm-diameter inlets are compared. Theoretical flow characteristics along the inlet surface were obtained from an axisymmetric potential flow and boundary layer analysis. The experimental data were obtained from low-speed tests of a high-bypass-ratio turbofan engine simulator. Comparisons between calculated internal surface pressure distributions and experimental data are presented for a free-system velocity of 45 m/sec and for incidence angles from 0 deg to 50 deg. Analysis of boundary layer separation on the inlet lip at incidence angle is the major emphasis of this report. Theoretical boundary layer shape factors, skin friction coefficients, and velocity profiles in the boundary layer are presented, along with the location of the transition region. Theoretical and experimental separation locations are also discussed.

Flow Batteries for Microfluidic Networks – Configuring An Electroosmotic Pump for Non-Terminal Positions

PubMed Central

He, Chiyang; Lu, Joann J.; Jia, Zhijian; Wang, Wei; Wang, Xiayan; Dasgupta, Purnendu K.; Liu, Shaorong

2011-01-01

A micropump provides flow and pressure for a lab-on-chip device, just as a battery supplies current and voltage for an electronic system. Numerous micropumps have been developed, but none is as versatile as a battery. One cannot easily insert a micropump into a nonterminal position of a fluidic line without affecting the rest of the fluidic system, one cannot simply connect several micropumps in series to enhance the pressure output, etc. In this work we develop a flow battery (or pressure power supply) to address this issue. A flow battery consists of a +EOP (in which the liquid flows in the same direction as the field gradient) and a −EOP (in which the liquid flows opposite to the electric field gradient), and the outlet of the +EOP is directly connected to the inlet of the −EOP. An external high voltage is applied to this outlet-inlet joint via a short gel-filled capillary that allows ions but not bulk liquid flow, while the +EOP’s inlet and the −EOP’s outlet (the flow battery’s inlet and outlet) are grounded. This flow battery can be deployed anywhere in a fluidic network without electrically affecting the rest of the system. Several flow batteries can be connected in series to enhance the pressure output to drive HPLC separations. In a fluidic system powered by flow batteries, a hydraulic Ohm’s law can be applied to analyze system pressures and flow rates. PMID:21375230

Supersonic Elliptical Ramp Inlet

NASA Technical Reports Server (NTRS)

Adamson, Eric E. (Inventor); Fink, Lawrence E. (Inventor); Fugal, Spencer R. (Inventor)

2016-01-01

A supersonic inlet includes a supersonic section including a cowl which is at least partially elliptical, a ramp disposed within the cowl, and a flow inlet disposed between the cowl and the ramp. The ramp may also be at least partially elliptical.

Performance of a high-work low aspect ratio turbine tested with a realistic inlet radial temperature profile

NASA Technical Reports Server (NTRS)

Stabe, R. G.; Whitney, W. J.; Moffitt, T. P.

1984-01-01

Experimental results are presented for a 0.767 scale model of the first stage of a two-stage turbine designed for a high by-pass ratio engine. The turbine was tested with both uniform inlet conditions and with an inlet radial temperature profile simulating engine conditions. The inlet temperature profile was essentially mixed-out in the rotor. There was also substantial underturning of the exit flow at the mean diameter. Both of these effects were attributed to strong secondary flows in the rotor blading. There were no significant differences in the stage performance with either inlet condition when differences in tip clearance were considered. Performance was very close to design intent in both cases. Previously announced in STAR as N84-24589

Blended Wing Body Systems Studies: Boundary Layer Ingestion Inlets With Active Flow Control

NASA Technical Reports Server (NTRS)

Geiselhart, Karl A. (Technical Monitor); Daggett, David L.; Kawai, Ron; Friedman, Doug

2003-01-01

A CFD analysis was performed on a Blended Wing Body (BWB) aircraft with advanced, turbofan engines analyzing various inlet configurations atop the aft end of the aircraft. The results are presented showing that the optimal design for best aircraft fuel efficiency would be a configuration with a partially buried engine, short offset diffuser using active flow control, and a D-shaped inlet duct that partially ingests the boundary layer air in flight. The CFD models showed that if active flow control technology can be satisfactorily developed, it might be able to control the inlet flow distortion to the engine fan face and reduce the powerplant performance losses to an acceptable level. The weight and surface area drag benefits of a partially submerged engine shows that it might offset the penalties of ingesting the low energy boundary layer air. The combined airplane performance of such a design might deliver approximately 5.5% better aircraft fuel efficiency over a conventionally designed, pod-mounted engine.

Development and Characterization Testing of an Air Pulsation Valve for a Pulse Detonation Engine Supersonic Parametric Inlet Test Section

NASA Technical Reports Server (NTRS)

Tornabene, Robert

2005-01-01

In pulse detonation engines, the potential exists for gas pulses from the combustor to travel upstream and adversely affect the inlet performance of the engine. In order to determine the effect of these high frequency pulses on the inlet performance, an air pulsation valve was developed to provide air pulses downstream of a supersonic parametric inlet test section. The purpose of this report is to document the design and characterization tests that were performed on a pulsation valve that was tested at the NASA Glenn Research Center 1x1 Supersonic Wind Tunnel (SWT) test facility. The high air flow pulsation valve design philosophy and analyses performed are discussed and characterization test results are presented. The pulsation valve model was devised based on the concept of using a free spinning ball valve driven from a variable speed electric motor to generate air flow pulses at preset frequencies. In order to deliver the proper flow rate, the flow port was contoured to maximize flow rate and minimize pressure drop. To obtain sharp pressure spikes the valve flow port was designed to be as narrow as possible to minimize port dwell time.

Electric fluid pump

DOEpatents

Van Dam, Jeremy Daniel; Turnquist, Norman Arnold; Raminosoa, Tsarafidy; Shah, Manoj Ramprasad; Shen, Xiaochun

2015-09-29

An electric machine is presented. The electric machine includes a hollow rotor; and a stator disposed within the hollow rotor, the stator defining a flow channel. The hollow rotor includes a first end portion defining a fluid inlet, a second end portion defining a fluid outlet; the fluid inlet, the fluid outlet, and the flow channel of the stator being configured to allow passage of a fluid from the fluid inlet to the fluid outlet via the flow channel; and wherein the hollow rotor is characterized by a largest cross-sectional area of hollow rotor, and wherein the flow channel is characterized by a smallest cross-sectional area of the flow channel, wherein the smallest cross-sectional area of the flow channel is at least about 25% of the largest cross-sectional area of the hollow rotor. An electric fluid pump and a power generation system are also presented.

Model investigation of inlet plenum flow straightening techniques for altitude test facility

NASA Technical Reports Server (NTRS)

Riddlebaugh, S. M.; Linke, H. G.

1976-01-01

An investigation was conducted to evaluate and improve the quality of the airflow to be supplied to the engine in altitude test chambers 3 and 4 of the Propulsion Systems Laboratory at the Lewis Research Center. One-twentieth-scale models of the inlet plenum chamber of the two test chambers were used in the investigation to minimize time and cost. It was possible to reduce the velocity spread in the inlet plenum from approximately 100 m/sec (330 ft/sec) to approximately 10 m/sec (30 ft/sec) through the combined use of flow diverters, multiple spaced screens, flow straighteners, and turning vanes.

Advanced prediction technique for the low speed aerodynamics of V/STOL aircraft. Volume 2: User's manual

NASA Technical Reports Server (NTRS)

Beatty, T. D.; Worthey, M. K.

1984-01-01

A computerized prediction method known as the Vought V/STOL Aircraft Propulsive Effects computer program (VAPE) for propulsive induced forces and moments in transition and Short TakeOff and Landing (STOL) flight is improved and evaluated. The VAPE program is capable of evaluating: (1) effects of relative wind about an aircraft, (2) effects of propulsive lift jet entrainment, vorticity and flow blockage, (3) effects of engine inlet flow on the aircraft flow field, (4) engine inlet forces and moments including inlet separation, (5) ground effects in the STOL region of flight, and (6) viscous effects on lifting surfaces.

System and method for improving performance of a fluid sensor for an internal combustion engine

DOEpatents

Kubinski, David [Canton, MI; Zawacki, Garry [Livonia, MI

2009-03-03

A system and method for improving sensor performance of an on-board vehicle sensor, such as an exhaust gas sensor, while sensing a predetermined substance in a fluid flowing through a pipe include a structure for extending into the pipe and having at least one inlet for receiving fluid flowing through the pipe and at least one outlet generally opposite the at least one inlet, wherein the structure redirects substantially all fluid flowing from the at least one inlet to the sensor to provide a representative sample of the fluid to the sensor before returning the fluid through the at least one outlet.

Detailed design of a quiet high flow fan

NASA Technical Reports Server (NTRS)

Soltau, J. D.; Orelup, M. J.; Beguhn, A. A.; Wiles, F. M.; Anderson, M. J.

1977-01-01

A single stage fan was designed to demonstrate the noise abatement properties of near-sonic inlet flow and long-chord stator vanes for the reduction of both upstream and downstream propagated fan source noise. It is designed to produce a pressure ratio of 1.653:1 with an adiabatic efficiency of 83.9%. The fan has a 508 mm inlet diameter with a hub/tip ratio of 0.426 and a design tip speed of 533.4 m/sec. The design inlet specific flow rate is 219.71 kg/sec sq m and there are 10 tandem stator vanes with a combined aspect ratio of 0.54.

Cooling system for a bearing of a turbine rotor

DOEpatents

Schmidt, Mark Christopher

2002-01-01

In a gas turbine, a bore tube assembly radially inwardly of an aft bearing conveys cooling steam to the buckets of the turbine and returns the cooling steam to a return. To cool the bearing and thermally insulate the bearing from the cooling steam paths, a radiation shield is spaced from the bore tube assembly by a dead air gap. Additionally, an air passageway is provided between the radiation shield and the inner surface of an aft shaft forming part of the rotor. Air is supplied from an inlet for flow along the passage and radially outwardly through bores in the aft shaft disk to cool the bearing and insulate it from transfer of heat from the cooling steam.

Unstart phenomena induced by flow choking in scramjet inlet-isolators

NASA Astrophysics Data System (ADS)

Im, Seong-kyun; Do, Hyungrok

2018-02-01

A review of recent research outcomes in downstream flow choking-driven unstart is presented. Unstart is a flow phenomenon at the inlet that severely reduces the air mass flow rate through the engine, causing a loss of thrust and considerable transient mechanical loading. Therefore, unstart in a scramjet engine crucially affects the design and the operation range of hypersonic vehicles. Downstream flow choking is known to be one of the major mechanisms inducing inlet unstart, as confirmed by recent scramjet-powered flight tests. The current paper examines recent research progress in identifying flow choking mechanisms that trigger unstart. Three different flow choking mechanisms are discussed: flow blockage, mass addition, and heat release from combustion reactions. Current research outcomes on the characteristic of unstarting flows, such as transient and quasi-steady motions, are reviewed for each flow choking mechanism. The characteristics of unstarted flows are described including Buzzing phenomena and oscillatory motions of unstarted shockwaves. Then, the state-of-the-art methods to predict, detect, and control unstart are presented. The review suggests that further investigations with high-enthalpy ground facilities will aid understanding of heat release-driven unstart.

Transient liquid-crystal technique used to produce high-resolution convective heat-transfer-coefficient maps

NASA Technical Reports Server (NTRS)

Hippensteele, Steven A.; Poinsatte, Philip E.

1993-01-01

In this transient technique the preheated isothermal model wall simulates the classic one-dimensional, semi-infinite wall heat transfer conduction problem. By knowing the temperature of the air flowing through the model, the initial temperature of the model wall, and the surface cooling rate measured at any location with time (using the fast-response liquid-crystal patterns recorded on video tape), the heat transfer coefficient can be calculated for the color isothermal pattern produced. Although the test was run transiently, the heat transfer coefficients are for the steady-state case. The upstream thermal boundary condition was considered to be isothermal. This transient liquid-crystal heat-transfer technique was used in a transient air tunnel in which a square-inlet, 3-to-1 exit transition duct was placed. The duct was preheated prior to allowing room temperature air to be suddenly drawn through it. The resulting isothermal contours on the duct surfaces were revealed using a surface coating of thermochromic liquid crystals that display distinctive colors at particular temperatures. A video record was made of the temperature and time data for all points on the duct surfaces during each test. The duct surfaces were uniformly heated using two heating systems: the first was an automatic temperature-controlled heater blanket completely surrounding the test duct like an oven, and the second was an internal hot-air loop through the inside of the test duct. The hot-air loop path was confined inside the test duct by insulated heat dams located at the inlet and exit ends of the test duct. A recirculating fan moved hot air into the duct inlet, through the duct, out of the duct exit, through the oven, and back to the duct inlet. The temperature nonuniformity of the test duct model wall was held very small. Test results are reported for two inlet Reynolds numbers of 200,000 and 1,150,000 (based on the square-inlet hydraulic diameter) and two free-stream turbulence intensities of about 1 percent, which is typical of wind tunnels, and up to 20 percent (using a grid), which is typical of real engine conditions.

Flow behavior in inlet guide vanes of radial turbines

NASA Technical Reports Server (NTRS)

Sokhey, J.; Tabakoff, W.; Hosny, W. M.

1975-01-01

Scroll flow is discussed. Streamline pattern and velocity distribution in the guide vanes are calculated. The blade surface temperature distribution is also determined. The effects of the blade shapes and the nozzle channel width on the velocity profiles at inlet to the guide vanes are investigated.

SST Technology Follow-on Program - Phase I, Performance Evaluation of an SST Noise Suppressor Nozzle System. Volume 1. Suppressed Mode.

DTIC Science & Technology

ACOUSTIC INSULATION, *TURBOJET EXHAUST NOZZLES, *JET ENGINE NOISE, REDUCTION, JET TRANSPORT AIRCRAFT, THRUST AUGMENTATION , SUPERSONIC NOZZLES, DUCT...INLETS, CONVERGENT DIVERGENT NOZZLES, SUBSONIC FLOW, SUPERSONIC FLOW, SUPPRESSORS, TURBOJET INLETS, BAFFLES, JET PUMPS, THRUST , DRAG, TEMPERATURE

Investigation of Aerodynamic and Icing Characteristics of a Flush Alternate Inlet Induction System Air Scoop

NASA Technical Reports Server (NTRS)

Lewis, James P.

1953-01-01

An investigation has been made in the NACA Lewis icing research tunnel to determine the aerodynamic and icing characteristics of a full-scale induction-system air-scoop assembly incorporating a flush alternate inlet. The flush inlet was located immediately downstream of the offset ram inlet and included a 180 deg reversal and a 90 deg elbow in the ducting between inlet and carburetor top deck. The model also had a preheat-air inlet. The investigation was made over a range of mass-air- flow ratios of 0 to 0.8, angles of attack of 0 and 4 deg airspeeds of 150 to 270 miles per hour, air temperatures of 0 and 25 F various liquid-water contents, and droplet sizes. The ram inlet gave good pressure recovery in both clear air and icing but rapid blockage of the top-deck screen occurred during icing. The flush alternate inlet had poor pressure recovery in both clear air and icing. The greatest decreases in the alternate-inlet pressure recovery were obtained at icing conditions of low air temperature and high liquid-water content. No serious screen icing was observed with the alternate inlet. Pressure and temperature distributions on the carburetor top deck were determined using the preheat-air supply with the preheat- and alternate-inlet doors in various positions. No screen icing occurred when the preheat-air system was operated in combination with alternate-inlet air flow.

Development of Wing Inlets

NASA Technical Reports Server (NTRS)

Racisz, Stanley F.

1946-01-01

Lift, drag, internal flow, and pressure distribution measurements were made on a low-drag airfoil incorporating various air inlet designs. Two leading-edge air inlets are developed which feature higher lift coefficients and critical Mach than the basic airfoil. Higher lift coefficients and critical speeds are obtained for leading half of these inlet sections but because of high suction pressures near exist, slightly lower critical speeds are obtained for the entire inlet section than the basic airfoil.

Background-Oriented Schlieren (BOS) for Scramjet Inlet-isolator Investigation

NASA Astrophysics Data System (ADS)

Che Idris, Azam; Rashdan Saad, Mohd; Hing Lo, Kin; Kontis, Konstantinos

2018-05-01

Background-oriented Schlieren (BOS) technique is a recently invented non-intrusive flow diagnostic method which has yet to be fully explored in its capabilities. In this paper, BOS technique has been applied for investigating the general flow field characteristics inside a generic scramjet inlet-isolator with Mach 5 flow. The difficulty in finding the delicate balance between measurement sensitivity and measurement area image focusing has been demonstrated. The differences between direct cross-correlation (DCC) and Fast Fourier Transform (FFT) raw data processing algorithm have also been demonstrated. As an exploratory study of BOS capability, this paper found that BOS is simple yet robust enough to be used to visualize complex flow in a scramjet inlet in hypersonic flow. However, in this case its quantitative data can be strongly affected by 3-dimensionality thus obscuring the density value with significant errors.

Luminescent Measurement Systems for the Investigation of a Scramjet Inlet-Isolator

PubMed Central

Idris, Azam Che; Saad, Mohd Rashdan; Zare-Behtash, Hossein; Kontis, Konstantinos

2014-01-01

Scramjets have become a main focus of study for many researchers, due to their application as propulsive devices in hypersonic flight. This entails a detailed understanding of the fluid mechanics involved to be able to design and operate these engines with maximum efficiency even at their off-design conditions. It is the objective of the present cold-flow investigation to study and analyse experimentally the mechanics of the fluid structures encountered within a generic scramjet inlet at M = 5. Traditionally, researchers have to rely on stream-thrust analysis, which requires the complex setup of a mass flow meter, a force balance and a heat transducer in order to measure inlet-isolator performance. Alternatively, the pitot rake could be positioned at inlet-isolator exit plane, but this method is intrusive to the flow, and the number of pitot tubes is limited by the model size constraint. Thus, this urgent need for a better flow diagnostics method is addressed in this paper. Pressure-sensitive paint (PSP) has been applied to investigate the flow characteristics on the compression ramp, isolator surface and isolator sidewall. Numerous shock-shock interactions, corner and shoulder separation regions, as well as shock trains were captured by the luminescent system. The performance of the scramjet inlet-isolator has been shown to improve when operated in a modest angle of attack. PMID:24721773

Method for Determining Optimum Injector Inlet Geometry

NASA Technical Reports Server (NTRS)

Myers, W. Neill (Inventor); Trinh, Huu P. (Inventor)

2015-01-01

A method for determining the optimum inlet geometry of a liquid rocket engine swirl injector includes obtaining a throttleable level phase value, volume flow rate, chamber pressure, liquid propellant density, inlet injector pressure, desired target spray angle and desired target optimum delta pressure value between an inlet and a chamber for a plurality of engine stages. The method calculates the tangential inlet area for each throttleable stage. The method also uses correlation between the tangential inlet areas and delta pressure values to calculate the spring displacement and variable inlet geometry of a liquid rocket engine swirl injector.

Direct numerical simulations of two-phase laminar jet flows with different cross-section injection geometries

NASA Astrophysics Data System (ADS)

Abdel-Hameed, H.; Bellan, J.

2002-10-01

Direct numerical simulations are performed of spatial, three-dimensional, laminar jets of different inlet geometric configurations for the purpose of quantifying the characteristics of the flows; both single-phase (SP) and two-phase (TP) free jets are considered. The TP jets consist of gas laden with liquid drops randomly injected at the inlet. Drop evaporation ensues both due to the gaseous flow being initially unvitiated by the vapor species corresponding to the liquid drops, and to drop heating as the initial drop temperature is lower than that of the carrier gas. The conservation equations for the TP flow include complete couplings of mass, momentum, and energy based on thermodynamically self-consistent specification of the vapor enthalpy, internal energy, and latent heat of vaporization. Inlet geometries investigated are circular, elliptic, rectangular, square, and triangular. The results focus both on the different spreading achieved according to the inlet geometry, as well as on the considerable change in the flow field due to the presence of the drops. The most important consequence of the drop interaction with the flow is the production of streamwise vorticity that alters entrainment and species mixing according to the inlet geometry. Similar to their SP equivalent, TP jets are shown to reach steady-state entrainment; examination of the flows at this time station shows that the potential cores of TP jets are shorter by an order of magnitude than their SP counterpart. Moreover, whereas the TP circular jet exhibits a symmetric entrainment pattern well past the streamwise location of the potential core, noncircular jets display at the same location strong departures from symmetry. Furthermore, the SP-jet phenomenon of axis switching is no longer present in TP jets. The distributions of drop-number density, liquid mass, and evaporated species are compared for different inlet cross sections and recommendations are made regarding the optimal choice for different applications.

Numerical simulation of groundwater flow at Puget Sound Naval Shipyard, Naval Base Kitsap, Bremerton, Washington

USGS Publications Warehouse

Jones, Joseph L.; Johnson, Kenneth H.; Frans, Lonna M.

2016-08-18

Information about groundwater-flow paths and locations where groundwater discharges at and near Puget Sound Naval Shipyard is necessary for understanding the potential migration of subsurface contaminants by groundwater at the shipyard. The design of some remediation alternatives would be aided by knowledge of whether groundwater flowing at specific locations beneath the shipyard will eventually discharge directly to Sinclair Inlet of Puget Sound, or if it will discharge to the drainage system of one of the six dry docks located in the shipyard. A 1997 numerical (finite difference) groundwater-flow model of the shipyard and surrounding area was constructed to help evaluate the potential for groundwater discharge to Puget Sound. That steady-state, multilayer numerical model with homogeneous hydraulic characteristics indicated that groundwater flowing beneath nearly all of the shipyard discharges to the dry-dock drainage systems, and only shallow groundwater flowing beneath the western end of the shipyard discharges directly to Sinclair Inlet.Updated information from a 2016 regional groundwater-flow model constructed for the greater Kitsap Peninsula was used to update the 1997 groundwater model of the Puget Sound Naval Shipyard. That information included a new interpretation of the hydrogeologic units underlying the area, as well as improved recharge estimates. Other updates to the 1997 model included finer discretization of the finite-difference model grid into more layers, rows, and columns, all with reduced dimensions. This updated Puget Sound Naval Shipyard model was calibrated to 2001–2005 measured water levels, and hydraulic characteristics of the model layers representing different hydrogeologic units were estimated with the aid of state-of-the-art parameter optimization techniques.The flow directions and discharge locations predicted by this updated model generally match the 1997 model despite refinements and other changes. In the updated model, most groundwater discharge recharged within the boundaries of the shipyard is to the dry docks; only at the western end of the shipyard does groundwater discharge directly to Puget Sound. Particle tracking for the existing long-term monitoring well network suggests that only a few wells intercept groundwater that originates as recharge within the shipyard boundary.

A zonal computational procedure adapted to the optimization of two-dimensional thrust augmentor inlets

NASA Technical Reports Server (NTRS)

Lund, T. S.; Tavella, D. A.; Roberts, L.

1985-01-01

A viscous-inviscid interaction methodology based on a zonal description of the flowfield is developed as a mean of predicting the performance of two-dimensional thrust augmenting ejectors. An inviscid zone comprising the irrotational flow about the device is patched together with a viscous zone containing the turbulent mixing flow. The inviscid region is computed by a higher order panel method, while an integral method is used for the description of the viscous part. A non-linear, constrained optimization study is undertaken for the design of the inlet region. In this study, the viscous-inviscid analysis is complemented with a boundary layer calculation to account for flow separation from the walls of the inlet region. The thrust-based Reynolds number as well as the free stream velocity are shown to be important parameters in the design of a thrust augmentor inlet.

Effect of inlet geometry on macrosegregation during the direct chill casting of 7050 alloy billets: experiments and computer modelling

NASA Astrophysics Data System (ADS)

Zhang, L.; Eskin, D. G.; Miroux, A.; Subroto, T.; Katgerman, L.

2012-07-01

Controlling macrosegregation is one of the major challenges in direct-chill (DC) casting of aluminium alloys. In this paper, the effect of the inlet geometry (which influences the melt distribution) on macrosegregation during the DC casting of 7050 alloy billets was studied experimentally and by using 2D computer modelling. The ALSIM model was used to determine the temperature and flow patterns during DC casting. The results from the computer simulations show that the sump profiles and flow patterns in the billet are strongly influenced by the melt flow distribution determined by the inlet geometry. These observations were correlated to the actual macrosegregation patterns found in the as-cast billets produced by having two different inlet geometries. The macrosegregation analysis presented here may assist in determining the critical parameters to consider for improving the casting of 7XXX aluminium alloys.

Statistical prediction of dynamic distortion of inlet flow using minimum dynamic measurement. An application to the Melick statistical method and inlet flow dynamic distortion prediction without RMS measurements

NASA Technical Reports Server (NTRS)

Schweikhard, W. G.; Chen, Y. S.

1986-01-01

The Melick method of inlet flow dynamic distortion prediction by statistical means is outlined. A hypothetic vortex model is used as the basis for the mathematical formulations. The main variables are identified by matching the theoretical total pressure rms ratio with the measured total pressure rms ratio. Data comparisons, using the HiMAT inlet test data set, indicate satisfactory prediction of the dynamic peak distortion for cases with boundary layer control device vortex generators. A method for the dynamic probe selection was developed. Validity of the probe selection criteria is demonstrated by comparing the reduced-probe predictions with the 40-probe predictions. It is indicated that the the number of dynamic probes can be reduced to as few as two and still retain good accuracy.

Blade row dynamic digital compressor program. Volume 1: J85 clean inlet flow and parallel compressor models

NASA Technical Reports Server (NTRS)

Tesch, W. A.; Steenken, W. G.

1976-01-01

The results are presented of a one-dimensional dynamic digital blade row compressor model study of a J85-13 engine operating with uniform and with circumferentially distorted inlet flow. Details of the geometry and the derived blade row characteristics used to simulate the clean inlet performance are given. A stability criterion based upon the self developing unsteady internal flows near surge provided an accurate determination of the clean inlet surge line. The basic model was modified to include an arbitrary extent multi-sector parallel compressor configuration for investigating 180 deg 1/rev total pressure, total temperature, and combined total pressure and total temperature distortions. The combined distortions included opposed, coincident, and 90 deg overlapped patterns. The predicted losses in surge pressure ratio matched the measured data trends at all speeds and gave accurate predictions at high corrected speeds where the slope of the speed lines approached the vertical.

Calculation of compressible flow in and about three-dimensional inlets with and without auxiliary inlets by a higher-order panel method

NASA Technical Reports Server (NTRS)

Hess, J. L.; Friedman, D. M.

1982-01-01

A three dimensional higher order panel method was specialized to the case of inlets with auxiliary inlets. The resulting program has a number of graphical input-output features to make it highly useful to the designer. The various aspects of the program are described instructions for its use are presented.

An experimental and computational investigation of flow in a radial inlet of an industrial pipeline centrifugal compressor

DOE Office of Scientific and Technical Information (OSTI.GOV)

Flathers, M.B.; Bache, G.E.; Rainsberger, R.

1996-04-01

The flow field of a complex three-dimensional radial inlet for an industrial pipeline centrifugal compressor has been experimentally determined on a half-scale model. Based on the experimental results, inlet guide vanes have been designed to correct pressure and swirl angle distribution deficiencies. The unvaned and vaned inlets are analyzed with a commercially available fully three-dimensional viscous Navier-Stokes code. Since experimental results were available prior to the numerical study, the unvaned analysis is considered a postdiction while the vaned analysis is considered a prediction. The computational results of the unvaned inlet have been compared to the previously obtained experimental results. Themore » experimental method utilized for the unvaned inlet is repeated for the vaned inlet and the data have been used to verify the computational results. The paper will discuss experimental, design, and computational procedures, grid generation, boundary conditions, and experimental versus computational methods. Agreement between experimental and computational results is very good, both in prediction and postdiction modes. The results of this investigation indicate that CFD offers a measurable advantage in design, schedule, and cost and can be applied to complex, three-dimensional radial inlets.« less

77 FR 30283 - Standards of Performance for New Stationary Sources, National Emission Standards for Hazardous...

Federal Register 2010, 2011, 2012, 2013, 2014

2012-05-22

... timeline and exceedances of the pressure, temperature, and oxygen and/or nitrogen concentration are... (AMP) that consists of monitoring the inlet scrubbing liquid temperature, flow rate, and acid content... request consisting of monitoring the inlet scrubbing liquid temperature and flow rate and identifying...

Analytical and Experimental Study of Flow Through an Axial Turbine Stage with a Nonuniform Inlet Radial Temperature Profile

NASA Technical Reports Server (NTRS)

Schwab, J. R.; Stabe, R. G.; Whitney, W. J.

1983-01-01

Results are presented for a typical nonuniform inlet radial temperature profile through an advanced single-stage axial turbine and compared with the results obtained for a uniform profile. Gas temperature rises of 40 K to 95 K are predicted at the hub and tip corners at the trailing edges of the pressure surfaces in both the stator and rotor due to convection of hot fluid from the mean by the secondary flow. The inlet temperature profile is shown to be mixed out at the rotor exit survey plane (2.3 axial chords downstream of the rotor trailing edge) in both the analysis and the experiment. The experimental rotor exit angle profile for the nonuniform inlet temperature profile indicates underturning at the tip caused by increased clearance. Severe underturning also occurs at the mean, both with and without the nonuniform inlet temperature profile. The inviscid rotational flow code used in the analysis fails to predict the underturning at the mean, which may be caused by viscous effects.

Analytical and experimental study of flow through an axial turbine stage with a nonuniform inlet radial temperature profile

NASA Technical Reports Server (NTRS)

Schwab, J. R.; Stabe, R. G.; Whitney, W. J.

1983-01-01

Results are presented for a typical nonuniform inlet radial temperature profile through an advanced single-stage axial turbine and compared with the results obtained for a uniform profile. Gas temperature rises of 40 K to 95 K are predicted at the hub and tip corners at the trailing edges of the pressure surfaces in both the stator and rotor due to convection of hot fluid from the mean by the secondary flow. The inlet temperature profile is shown to be mixed out at the rotor exit survey plane (2.3 axial chords downstream of the rotor trailing edge) in both the analysis and the experiment. The experimental rotor exit angle profile for the nonuniform inlet temperature profile indicates underturning at the tip caused by increased clearance. Severe underturning also occurs at the mean, both with and without the nonuniform inlet temperature profile. The inviscid rotational flow code used in the analysis fails to predict the underturning at the mean, which may be caused by viscous effects. Previously announced in STAR as N83-27958

Preliminary Investigation of a New Type of Supersonic Inlet

NASA Technical Reports Server (NTRS)

Ferri, Antonio; Nucci, Louis M

1946-01-01

A supersonic inlet with supersonic deceleration of the flow entirely outside of the inlet is considered. A particular arrangement with fixed geometry having a central body with a circular annular intake is analyzed, and it is shown theoretically that this arrangement gives high pressure recovery for a large range of Mach number and mass flow and therefore is practical for use on supersonic airplanes and missiles. For some Mach numbers the drag coefficient for this type of inlet is larger than the drag coefficient for the type of inlet with supersonic compression entirely inside, but the pressure recovery is larger for all flight conditions. The differences in drag can be eliminated for the design Mach number. Experimental results confirm the results of the theoretical analysis and show that pressure recoveries of 95 percent for Mach numbers of 1.33 and 1.52, 92 percent for a Mach number of 1.72, and 86 percent for a Mach number oof 2.10 are possible with the configurations considered. If the mass flow decreases, the total drag coefficient increases gradually and the pressure recovery does not change appreciably.

Solar Cell Modules With Improved Backskin

DOEpatents

Chevrefils, Andre; Grigore, Daniel Gheorghe

2001-01-23

The present invention relates to gas turbines and more particularly to a device for controlling the flow of cooling air through a flowpath in a turbine blade. The device can be inserted in the inlet opening of the blade flowpath and be retained therein. The device comprises a plug member for adjusting the flow of cooling air through the flowpath. The plug member comprises a retaining portion for retaining the plug member at the inlet opening of the flowpath and a blocking portion inserted within the flowpath for reducing the cross-sectional area of the inlet opening. Such a device is inexpensive and can be easily inserted in the inlet opening of a blade flowpath and retained therein.

A Framework for Simulating Turbine-Based Combined-Cycle Inlet Mode-Transition

NASA Technical Reports Server (NTRS)

Le, Dzu K.; Vrnak, Daniel R.; Slater, John W.; Hessel, Emil O.

2012-01-01

A simulation framework based on the Memory-Mapped-Files technique was created to operate multiple numerical processes in locked time-steps and send I/O data synchronously across to one-another to simulate system-dynamics. This simulation scheme is currently used to study the complex interactions between inlet flow-dynamics, variable-geometry actuation mechanisms, and flow-controls in the transition from the supersonic to hypersonic conditions and vice-versa. A study of Mode-Transition Control for a high-speed inlet wind-tunnel model with this MMF-based framework is presented to illustrate this scheme and demonstrate its usefulness in simulating supersonic and hypersonic inlet dynamics and controls or other types of complex systems.

Rotor Noise due to Blade-Turbulence Interaction.

NASA Astrophysics Data System (ADS)

Ishimaru, Kiyoto

The time-averaged intensity density function of the acoustic radiation from rotating blades is derived by replacing blades with rotating dipoles. This derivation is done under the following turbulent inflow conditions: turbulent ingestion with no inlet strut wakes, inflow turbulence elongation and contraction with no inlet strut wakes, and inlet strut wakes. Dimensional analysis reveals two non-dimensional parameters which play important roles in generating the blade-passing frequency tone and its multiples. The elongation and contraction of inflow turbulence has a strong effect on the generation of the blade-passing frequency tone and its multiples. Increasing the number of rotor blades widens the peak at the blade-passing frequency and its multiples. Increasing the rotational speed widens the peak under the condition that the non-dimensional parameter involving the rotational speed is fixed. The number of struts and blades should be chosen so that (the least common multiple of them)(.)(rotational speed) is in the cutoff range of Sears' function, in order to minimize the effect of the mean flow deficit on the time averaged intensity density function. The acoustic intensity density function is not necessarily stationary even if the inflow turbulence is homogeneous and isotropic. The time variation of the propagation path due to the rotation should be considered in the computation of the intensity density function; for instance, in the present rotor specification, the rotor radius is about 0.3 m and the rotational speed Mach number is about 0.2.

DESIGN ANALYSIS OF RADIAL INFLOW TURBINES

NASA Technical Reports Server (NTRS)

Glassman, A. J.

1994-01-01

This program performs a velocity-diagram analysis required for determining geometry and estimating performance for radial-inflow turbines. Input design requirements are power, mass flow rate, inlet temperature and pressure, and rotative rate. The design variables include stator-exit angle, rotor-exit-tip to rotor-inlet radius ratio, rotor-exit-hub to tip radius ratio, and the magnitude and radial distribution of rotor-exit tangential velocity. The program output includes diameters, total and static efficiences, all absolute and relative temperatures, pressures, and velocities, and flow angles at stator inlet, stator exit, rotor inlet, and rotor exit. Losses accounted for in this program by the internal loss model are three-dimensional (profile plus end wall) viscous losses in the stator and the rotor, the disk-friction loss on the back side of the rotor, the loss due to the clearance between the rotor tip and the outer casing, and the exit velocity loss. The flow analysis is one-dimensional at the stator inlet, stator exit, and rotor inlet, each of these calculation stations being at a constant radius. At the rotor exit where there is a variation in flow-field radius, an axisymmetric two-dimensional analysis is made using constant height sectors. Simple radial equilibrium is used to establish the static pressure gradient at the rotor exit. This program is written in FORTRAN V and has been implemented on a UNIVAC 1100 series computer with a memory requirement of approximately 22K of 36 bit words.

Inward-Turning Streamline-Traced Inlet Design Method for Low-Boom, Low-Drag Applications

NASA Technical Reports Server (NTRS)

Otto, Samuel; Trefny, Charles J.; Slater, John W.

2015-01-01

A new design method for inward-turning, streamline-traced inlets is presented. Resulting designs are intended for moderate supersonic, low-drag, low-boom applications such as that required for NASA's proposed low-boom flight demonstration aircraft. A critical feature of these designs is the internal cowl lip angle that allows for little or no flow turning on the outer nacelle. Present methods using conical-flow Busemann parent flowfields have simply truncated, or otherwise modified the stream-traced contours to include this internal cowl angle. Such modifications disrupt the parent flowfield, reducing inlet performance and flow uniformity. The method presented herein merges a conical flowfield that includes a leading shock with a truncated Busemann flowfield in a manner that minimizes unwanted interactions. A leading internal cowl angle is now inherent in the parent flowfield, and inlet contours traced from this flowfield retain its high performance and good flow uniformity. CFD analysis of a candidate inlet design is presented that verifies the design technique, and reveals a starting issue with the basic geometry. A minor modification to the cowl lip region is shown to eliminate this phenomenon, thereby allowing starting and smooth transition to sub-critical operation as back-pressure is increased. An inlet critical-point total pressure recovery of 96 is achieved based on CFD results for a Mach 1.7 freestream design. Correction for boundary-layer displacement thickness, and sizing for a given engine airflow requirement are also discussed.

Confined turbulent swirling recirculating flow predictions. Ph.D. Thesis. Final Report

NASA Technical Reports Server (NTRS)

Abujelala, M. T.; Lilley, D. G.

1985-01-01

The capability and the accuracy of the STARPIC computer code in predicting confined turbulent swirling recirculating flows is presented. Inlet flow boundary conditions were demonstrated to be extremely important in simulating a flowfield via numerical calculations. The degree of swirl strength and expansion ratio have strong effects on the characteristics of swirling flow. In a nonswirling flow, a large corner recirculation zone exists in the flowfield with an expansion ratio greater than one. However, as the degree of inlet swirl increases, the size of this zone decreases and a central recirculation zone appears near the inlet. Generally, the size of the central zone increased with swirl strength and expansion ratio. Neither the standard k-epsilon turbulence mode nor its previous extensions show effective capability for predicting confined turbulent swirling recirculating flows. However, either reduced optimum values of three parameters in the mode or the empirical C sub mu formulation obtained via careful analysis of available turbulence measurements, can provide more acceptable accuracy in the prediction of these swirling flows.

Aerosol mobility size spectrometer

DOEpatents

Wang, Jian; Kulkarni, Pramod

2007-11-20

A device for measuring aerosol size distribution within a sample containing aerosol particles. The device generally includes a spectrometer housing defining an interior chamber and a camera for recording aerosol size streams exiting the chamber. The housing includes an inlet for introducing a flow medium into the chamber in a flow direction, an aerosol injection port adjacent the inlet for introducing a charged aerosol sample into the chamber, a separation section for applying an electric field to the aerosol sample across the flow direction and an outlet opposite the inlet. In the separation section, the aerosol sample becomes entrained in the flow medium and the aerosol particles within the aerosol sample are separated by size into a plurality of aerosol flow streams under the influence of the electric field. The camera is disposed adjacent the housing outlet for optically detecting a relative position of at least one aerosol flow stream exiting the outlet and for optically detecting the number of aerosol particles within the at least one aerosol flow stream.

Flow control in axial fan inlet guide vanes by synthetic jets

NASA Astrophysics Data System (ADS)

Cyrus, V.; Trávníček, Z.; Wurst, P.; Kordík, J.

2013-04-01

Tested high pressure axial flow fan with hub/tip ratio of 0.70 and external diameter of 600 mm consisted of inlet guide vanes (IGV), rotor and stator blade rows. Fan peripheral velocity was 47 m/s. Air volume flow rate was changed by turning of rear part of the inlet guide vanes. At turning of 20 deg the flow was separated on the IGV profiles. The synthetic jets were introduced through radial holes in machine casing in the location before flow separation origin. Synthetic jet actuator was designed with the use of a speaker by UT AVCR. Its membrane had diameter of 63 mm. Excitation frequency was chosen in the range of 500 Hz - 700 Hz. Synthetic jets favourably influenced separated flow on the vane profiles in the distance of (5 - 12) mm from the casing surface. The reduction of flow separation area caused in the region near the casing the decrease of the profile loss coefficient approximately by 20%.

Design and Analysis Tools for Supersonic Inlets

NASA Technical Reports Server (NTRS)

Slater, John W.; Folk, Thomas C.

2009-01-01

Computational tools are being developed for the design and analysis of supersonic inlets. The objective is to update existing tools and provide design and low-order aerodynamic analysis capability for advanced inlet concepts. The Inlet Tools effort includes aspects of creating an electronic database of inlet design information, a document describing inlet design and analysis methods, a geometry model for describing the shape of inlets, and computer tools that implement the geometry model and methods. The geometry model has a set of basic inlet shapes that include pitot, two-dimensional, axisymmetric, and stream-traced inlet shapes. The inlet model divides the inlet flow field into parts that facilitate the design and analysis methods. The inlet geometry model constructs the inlet surfaces through the generation and transformation of planar entities based on key inlet design factors. Future efforts will focus on developing the inlet geometry model, the inlet design and analysis methods, a Fortran 95 code to implement the model and methods. Other computational platforms, such as Java, will also be explored.

Investigation of Blade Angle of an Open Cross-Flow Runner

NASA Astrophysics Data System (ADS)

Katayama, Yusuke; Iio, Shouichiro; Veerapun, Salisa; Uchiyama, Tomomi

2015-04-01

The aim of this study was to develop a nano-hydraulic turbine utilizing drop structure in irrigation channels or industrial waterways. This study was focused on an open-type cross-flow turbine without any attached equipment for cost reduction and easy maintenance. In this study, the authors used an artificial indoor waterfall as lab model. Test runner which is a simple structure of 20 circular arc-shaped blades sandwiched by two circular plates was used The optimum inlet blade angle and the relationship between the power performance and the flow rate approaching theoretically and experimentally were investigated. As a result, the optimum inlet blade angle due to the flow rate was changed. Additionally, allocation rate of power output in 1st stage and 2nd stage is changed by the blade inlet angle.

Computational Fluid Dynamics (CFD) study of the 4th generation prototype of a continuous flow Ventricular Assist Device (VAD).

PubMed

Song, Xinwei; Wood, Houston G; Olsen, Don

2004-04-01

The continuous flow ventricular assist device (VAD) is a miniature centrifugal pump, fully suspended by magnetic bearings, which is being developed for implantation in humans. The CF4 model is the first actual prototype of the final design product. The overall performances of blood flow in CF4 have been simulated using computational fluid dynamics (CFD) software: CFX, which is commercially available from ANSYS Inc. The flow regions modeled in CF4 include the inlet elbow, the five-blade impeller, the clearance gap below the impeller, and the exit volute. According to different needs from patients, a wide range of flow rates and revolutions per minute (RPM) have been studied. The flow rate-pressure curves are given. The streamlines in the flow field are drawn to detect stagnation points and vortices that could lead to thrombosis. The stress is calculated in the fluid field to estimate potential hemolysis. The stress is elevated to the decreased size of the blood flow paths through the smaller pump, but is still within the safe range. The thermal study on the pump, the blood and the surrounding tissue shows the temperature rise due to magnetoelectric heat sources and thermal dissipation is insignificant. CFD simulation proved valuable to demonstrate and to improve the performance of fluid flow in the design of a small size pump.

Coupled Analysis of an Inlet and Fan for a Quiet Supersonic Aircraft

NASA Technical Reports Server (NTRS)

Chima, Rodrick V.; Conners, Timothy R.; Wayman, Thomas R.

2009-01-01

A computational analysis of a Gulfstream isentropic external compression supersonic inlet coupled to a Rolls-Royce fan was completed. The inlet was designed for a small, low sonic boom supersonic vehicle with a design cruise condition of M = 1.6 at 45,000 feet. The inlet design included an annular bypass duct that routed flow subsonically around an engine-mounted gearbox and diverted flow with high shock losses away from the fan tip. Two Reynolds-averaged Navier-Stokes codes were used for the analysis: an axisymmetric code called AVCS for the inlet and a 3-D code called SWIFT for the fan. The codes were coupled at a mixing plane boundary using a separate code for data exchange. The codes were used to determine the performance of the inlet/fan system at the design point and to predict the performance and operability of the system over the flight profile. At the design point the core inlet had a recovery of 96 percent, and the fan operated near its peak efficiency and pressure ratio. A large hub radial distortion generated in the inlet was not eliminated by the fan and could pose a challenge for subsequent booster stages. The system operated stably at all points along the flight profile. Reduced stall margin was seen at low altitude and Mach number where flow separated on the interior lips of the cowl and bypass ducts. The coupled analysis gave consistent solutions at all points on the flight profile that would be difficult or impossible to predict by analysis of isolated components.

Coupled Analysis of an Inlet and Fan for a Quiet Supersonic Jet

NASA Technical Reports Server (NTRS)

Chima, Rodrick V.; Conners, Timothy R.; Wayman, Thomas R.

2010-01-01

A computational analysis of a Gulfstream isentropic external compression supersonic inlet coupled to a Rolls-Royce fan has been completed. The inlet was designed for a small, low sonic boom supersonic vehicle with a design cruise condition of M = 1.6 at 45,000 ft. The inlet design included an annular bypass duct that routed flow subsonically around an engine-mounted gearbox and diverted flow with high shock losses away from the fan tip. Two Reynolds-averaged Navier-Stokes codes were used for the analysis: an axisymmetric code called AVCS for the inlet and a three dimensional (3-D) code called SWIFT for the fan. The codes were coupled at a mixing plane boundary using a separate code for data exchange. The codes were used to determine the performance of the inlet/fan system at the design point and to predict the performance and operability of the system over the flight profile. At the design point the core inlet had a recovery of 96 percent, and the fan operated near its peak efficiency and pressure ratio. A large hub radial distortion generated in the inlet was not eliminated by the fan and could pose a challenge for subsequent booster stages. The system operated stably at all points along the flight profile. Reduced stall margin was seen at low altitude and Mach number where flow separated on the interior lips of the cowl and bypass ducts. The coupled analysis gave consistent solutions at all points on the flight profile that would be difficult or impossible to predict by analysis of isolated components.

Experimental study of the influence of flow passage subtle variation on mixed-flow pump performance

NASA Astrophysics Data System (ADS)

Bing, Hao; Cao, Shuliang

2014-05-01

In the mixed-flow pump design, the shape of the flow passage can directly affect the flow capacity and the internal flow, thus influencing hydraulic performance, cavitation performance and operation stability of the mixed-flow pump. However, there is currently a lack of experimental research on the influence mechanism. Therefore, in order to analyze the effects of subtle variations of the flow passage on the mixed-flow pump performance, the frustum cone surface of the end part of inlet contraction flow passage of the mixed-flow pump is processed into a cylindrical surface and a test rig is built to carry out the hydraulic performance experiment. In this experiment, parameters, such as the head, the efficiency, and the shaft power, are measured, and the pressure fluctuation and the noise signal are also collected. The research results suggest that after processing the inlet flow passage, the head of the mixed-flow pump significantly goes down; the best efficiency of the mixed-flow pump drops by approximately 1.5%, the efficiency decreases more significantly under the large flow rate; the shaft power slightly increases under the large flow rate, slightly decreases under the small flow rate. In addition, the pressure fluctuation amplitudes on both the impeller inlet and the diffuser outlet increase significantly with more drastic pressure fluctuations and significantly lower stability of the internal flow of the mixed-flow pump. At the same time, the noise dramatically increases. Overall speaking, the subtle variation of the inlet flow passage leads to a significant change of the mixed-flow pump performance, thus suggesting a special attention to the optimization of flow passage. This paper investigates the influence of the flow passage variation on the mixed-flow pump performance by experiment, which will benefit the optimal design of the flow passage of the mixed-flow pump.

Comparison Between Numerically Simulated and Experimentally Measured Flowfield Quantities Behind a Pulsejet

NASA Technical Reports Server (NTRS)

Geng, Tao; Paxson, Daniel E.; Zheng, Fei; Kuznetsov, Andrey V.; Roberts, William L.

2008-01-01

Pulsed combustion is receiving renewed interest as a potential route to higher performance in air breathing propulsion systems. Pulsejets offer a simple experimental device with which to study unsteady combustion phenomena and validate simulations. Previous computational fluid dynamic (CFD) simulation work focused primarily on the pulsejet combustion and exhaust processes. This paper describes a new inlet sub-model which simulates the fluidic and mechanical operation of a valved pulsejet head. The governing equations for this sub-model are described. Sub-model validation is provided through comparisons of simulated and experimentally measured reed valve motion, and time averaged inlet mass flow rate. The updated pulsejet simulation, with the inlet sub-model implemented, is validated through comparison with experimentally measured combustion chamber pressure, inlet mass flow rate, operational frequency, and thrust. Additionally, the simulated pulsejet exhaust flowfield, which is dominated by a starting vortex ring, is compared with particle imaging velocimetry (PIV) measurements on the bases of velocity, vorticity, and vortex location. The results show good agreement between simulated and experimental data. The inlet sub-model is shown to be critical for the successful modeling of pulsejet operation. This sub-model correctly predicts both the inlet mass flow rate and its phase relationship with the combustion chamber pressure. As a result, the predicted pulsejet thrust agrees very well with experimental data.

Experimental investigation of the radiation of sound from an unflanged duct and a bellmouth, including the flow effect

NASA Technical Reports Server (NTRS)

Ville, J. M.; Silcox, R. J.

1980-01-01

The radiation of sound from an inlet as a function of flow velocity, frequency, duct mode structure, and inlet geometry was examined by using a spinning mode synthesizer to insure a given space-time structure inside the duct. Measurements of the radiation pattern (amplitude and phase) and of the pressure reflection coefficient were obtained over an azimuthal wave number range of 0 to 6 and a frequency range up to 5000 Hz for an unflanged duct and a bellmouth. The measured radiated field and pressure reflection coefficient without flow for the unflanged duct agree reasonably well with theory. The influence of the inlet contour appears to be very drastic near the cut-on frequency of a mode and reasonable agreement is found between the bellmouth pressure reflection coefficient and a infinite hyperboloidal inlet theory. It is also shown that the flow has a weak effect on the amplitude of the directivity factor but significantly shifts the directivity factor phase. The influence of the flow on the modulus of the pressure reflection coefficient is found to be well described by a theoretical prediction.

System and method for continuous solids slurry depressurization

DOE Office of Scientific and Technical Information (OSTI.GOV)

Leininger, Thomas Frederick; Steele, Raymond Douglas; Cordes, Stephen Michael

A system includes a first pump having a first outlet and a first inlet, and a controller. The first pump is configured to continuously receive a flow of a slurry into the first outlet at a first pressure and to continuously discharge the flow of the slurry from the first inlet at a second pressure less than the first pressure. The controller is configured to control a first speed of the first pump against the flow of the slurry based at least in part on the first pressure, wherein the first speed of the first pump is configured to resistmore » a backflow of the slurry from the first outlet to the first inlet.« less

Effects of installation caused flow distortion on noise from a fan designed for turbofan engines

NASA Technical Reports Server (NTRS)

Povinelli, F. P.; Dittmar, J. H.; Woodward, R. P.

1972-01-01

Far-field noise measurements were taken for three different installations of essentially the same fan. The installation with the most uniform inlet flow resulted in fan-blade-passage tone sound pressure levels more than 10 dB lower than the installation with more nonuniform inflow. Perceived noise levels were computed for the various installations and compared. Some measurements of inlet flow distortion were made and used in a blade-passage noise generation theory to predict the effects of distortion on noise. Good agreement was obtained between the prediction and the measured effect. Possible origins of the distortion were identified by observation of tuft action in the vicinity of the inlet.

An investigation of bleed configurations and their effect on shock wave/boundary layer interactions

NASA Technical Reports Server (NTRS)

Hamed, Awatef

1995-01-01

The design of high efficiency supersonic inlets is a complex task involving the optimization of a number of performance parameters such as pressure recovery, spillage, drag, and exit distortion profile, over the flight Mach number range. Computational techniques must be capable of accurately simulating the physics of shock/boundary layer interactions, secondary corner flows, flow separation, and bleed if they are to be useful in the design. In particular, bleed and flow separation, play an important role in inlet unstart, and the associated pressure oscillations. Numerical simulations were conducted to investigate some of the basic physical phenomena associated with bleed in oblique shock wave boundary layer interactions that affect the inlet performance.

Investigation of flow fields within large scale hypersonic inlet models

NASA Technical Reports Server (NTRS)

Gnos, A. V.; Watson, E. C.; Seebaugh, W. R.; Sanator, R. J.; Decarlo, J. P.

1973-01-01

Analytical and experimental investigations were conducted to determine the internal flow characteristics in model passages representative of hypersonic inlets for use at Mach numbers to about 12. The passages were large enough to permit measurements to be made in both the core flow and boundary layers. The analytical techniques for designing the internal contours and predicting the internal flow-field development accounted for coupling between the boundary layers and inviscid flow fields by means of a displacement-thickness correction. Three large-scale inlet models, each having a different internal compression ratio, were designed to provide high internal performance with an approximately uniform static-pressure distribution at the throat station. The models were tested in the Ames 3.5-Foot Hypersonic Wind Tunnel at a nominal free-stream Mach number of 7.4 and a unit free-stream Reynolds number of 8.86 X one million per meter.

Optimal Micro-Vane Flow Control for Compact Air Vehicle Inlets

NASA Technical Reports Server (NTRS)

Anderson, Bernhard H.; Miller, Daniel N.; Addington, Gregory A.; Agrell, Johan

2004-01-01

The purpose of this study on micro-vane secondary flow control is to demonstrate the viability and economy of Response Surface Methodology (RSM) to optimally design micro-vane secondary flow control arrays, and to establish that the aeromechanical effects of engine face distortion can also be included in the design and optimization process. These statistical design concepts were used to investigate the design characteristics of "low unit strength" micro-effector arrays. "Low unit strength" micro-effectors are micro-vanes set at very low angles-of-incidence with very long chord lengths. They were designed to influence the near wall inlet flow over an extended streamwise distance, and their advantage lies in low total pressure loss and high effectiveness in managing engine face distortion. Therefore, this report examines optimal micro-vane secondary flow control array designs for compact inlets through a Response Surface Methodology.

Navier-Stokes simulation of external/internal transonic flow on the forebody/inlet of the AV-8B Harrier II

NASA Technical Reports Server (NTRS)

Mysko, Stephen J.; Chyu, Wei J.; Stortz, Michael W.; Chow, Chuen-Yen

1993-01-01

In this work, the computation of combined external/internal transonic flow on the complex forebody/inlet configuration of the AV-8B Harrier II is performed. The actual aircraft has been measured and its surface and surrounding domain, in which the fuselage and inlet have a common wall, have been described using structured grids. The 'thin-layer' Navier-Stokes equations were used to model the flow along with the Chimera embedded multi-block technique. A fully conservative, alternating direction implicit (ADI), approximately factored, partially fluxsplit algorithm was employed to perform the computation. Comparisons to some experimental wind tunnel data yielded good agreement for flow at zero incidence and angle of attack. The aim of this paper is to provide a methodology or computational tool for the numerical solution of complex external/internal flows.

Velocity and pressure measurements in guide vane clearance gap of a low specific speed Francis turbine

NASA Astrophysics Data System (ADS)

Thapa, B. S.; Dahlhaug, O. G.; Thapa, B.

2016-11-01

In Francis turbine, a small clearance gap between the guide vanes and the cover plates is usually required to pivot guide vanes as a part of governing system. Deflection of cover plates and erosion of mating surfaces causes this gap to increase from its design value. The clearance gap induces the secondary flow in the distributor system. This effects the main flow at the runner inlet, which causes losses in efficiency and instability. A guide vane cascade of a low specific speed Francis turbine has been developed for experimental investigations. The test setup is able to produce similar velocity distributions at the runner inlet as that of a reference prototype turbine. The setup is designed for particle image velocimetry (PIV) measurements from the position of stay vane outlet to the position of runner inlet. In this study, velocity and pressure measurements are conducted with 2 mm clearance gap on one side of guide vane. Leakage flow is observed and measured together with pressure measurements. It is concluded that the leakage flow behaves as a jet and mixes with the main flow in cross-wise direction and forms a vortex filament. This causes non-uniform inlet flow conditions at runner blades.

Aerodynamic Effects of Turbulence Intensity on a Variable-Speed Power-Turbine Blade with Large Incidence and Reynolds Number Variations

NASA Technical Reports Server (NTRS)

Flegel, Ashlie Brynn; Giel, Paul W.; Welch, Gerard E.

2014-01-01

The effects of inlet turbulence intensity on the aerodynamic performance of a variable speed power turbine blade are examined over large incidence and Reynolds number ranges. Both high and low turbulence studies were conducted in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. The purpose of the low inlet turbulence study was to examine the transitional flow effects that are anticipated at cruise Reynolds numbers. The high turbulence study extends this to LPT-relevant turbulence levels while perhaps sacrificing transitional flow effects. Downstream total pressure and exit angle data were acquired for ten incidence angles ranging from +15.8 to 51.0. For each incidence angle, data were obtained at five flow conditions with the exit Reynolds number ranging from 2.12105 to 2.12106 and at a design exit Mach number of 0.72. In order to achieve the lowest Reynolds number, the exit Mach number was reduced to 0.35 due to facility constraints. The inlet turbulence intensity, Tu, was measured using a single-wire hotwire located 0.415 axial-chord upstream of the blade row. The inlet turbulence levels ranged from 0.25 - 0.4 for the low Tu tests and 8- 15 for the high Tu study. Tu measurements were also made farther upstream so that turbulence decay rates could be calculated as needed for computational inlet boundary conditions. Downstream flow field measurements were obtained using a pneumatic five-hole pitchyaw probe located in a survey plane 7 axial chord aft of the blade trailing edge and covering three blade passages. Blade and endwall static pressures were acquired for each flow condition as well. The blade loading data show that the suction surface separation that was evident at many of the low Tu conditions has been eliminated. At the extreme positive and negative incidence angles, the data show substantial differences in the exit flow field. These differences are attributable to both the higher inlet Tu directly and to the thinner inlet endwall boundary layer that the turbulence grid imposes.

Theoretical analysis of incompressible flow through a radial-inlet centrifugal impeller at various weight flows

NASA Technical Reports Server (NTRS)

Kramer, James J; Prian, Vasily D; Wu, Chung-Hua

1956-01-01

A method for the solution of the incompressible nonviscous flow through a centrifugal impeller, including the inlet region, is presented. Several numerical solutions are obtained for four weight flows through an impeller at one operating speed. These solutions are refined in the leading-edge region. The results are presented in a series of figures showing streamlines and relative velocity contours. A comparison is made with the results obtained by using a rapid approximate method of analysis.

Numerical simulation of flow and melting characteristics of seawater-ice crystals two-phase flow in inlet straight pipe of shell and tube heat exchanger of polar ship

NASA Astrophysics Data System (ADS)

Xu, Li; Huang, Chang-Xu; Huang, Zhen-Fei; Sun, Qiang; Li, Jie

2018-05-01

The ice crystal particles are easy to enter into the seawater cooling system of polar ship together with seawater when it sails in the Arctic. They are easy to accumulate in the pipeline, causing serious blockage of the cooling pipe. In this study, the flow and melting characteristics of ice particles-seawater two-phase flow in inlet straight pipe of shell-and-tube heat exchanger were numerically simulated by using Eulerian-Eulerian two-fluid model coupled with the interphase heat and mass transfer model. The influences of inlet ice packing factor, ice crystal particle diameter, and inlet velocity on the distribution and melting characteristics of ice crystals were investigated. The degree of asymmetry of the distribution of ice crystals in the cross section decreases gradually when the IPF changes from 5 to 15%. The volume fractions of ice crystals near the top of the outlet cross section are 19.59, 19.51, and 22.24% respectively for ice packing factor of 5, 10 and 15%. When the particle diameter is 0.5 mm, the ice crystals are gradually stratified during the flow process. With particle diameters of 1.0 and 2.0 mm, the region with the highest volume fraction of ice crystals is a small circle and the contours in the cloud map are compact. The greater the inlet flow velocity, the less stratified the ice crystals and the more obvious the turbulence on the outlet cross section. The average volume fraction of ice crystals along the flow direction is firstly rapidly reduced and then stabilized after 300 mm.

Effects of inlet flow field conditions on the performance of centrifugal compressor diffusers: Part 2 -- Straight-channel diffuser

DOE Office of Scientific and Technical Information (OSTI.GOV)

Deniz, S.; Greitzer, E.M.; Cumpsty, N.A.

2000-01-01

This is Part 2 of an examination of the influence of inlet flow conditions on the performance and operating range of centrifugal compressor vaned diffusers. The paper describes tests of a straight-channel type diffuser, sometimes called a wedge-vane diffuser, and compares the results with those from the discrete-passage diffusers described in Part 1. Effects of diffuser inlet Mach number, flow angle, blockage, and axial flow nonuniformity on diffuser pressure recovery and operating range are addressed. The straight-channel diffuser investigated has 30 vanes and was designed for the same aerodynamic duty as the discrete-passage diffuser described in Part 1. The rangesmore » of the overall pressure recovery coefficients were 0.50--0.78 for the straight-channel diffuser and 0.50--0.70 for the discrete-passage diffuser, except when the diffuser was choked. In other words, the maximum pressure recovery of the straight-channel diffuser was found to be roughly 10% higher than that of the discrete-passage diffuser investigated. The two types of diffuser showed similar behavior regarding the dependence of pressure recovery on diffuser inlet flow angle and the insensitivity of the performance to inlet flow field axial distortion and Mach number. The operating range of the straight-channel diffuser, as for the discrete-passage diffusers, was limited by the onset of rotating stall at a fixed momentum-averaged flow angle into the diffuser, which was for the straight-channel diffuser, {alpha}{sub crit} = 70 {+-} 0.5 deg. The background, nomenclature, and description of the facility and method are all given in Part 1.« less

Comparison of numerical and experimental results of the flow in the U9 Kaplan turbine model

NASA Astrophysics Data System (ADS)

Petit, O.; Mulu, B.; Nilsson, H.; Cervantes, M.

2010-08-01

The present work compares simulations made using the OpenFOAM CFD code with experimental measurements of the flow in the U9 Kaplan turbine model. Comparisons of the velocity profiles in the spiral casing and in the draft tube are presented. The U9 Kaplan turbine prototype located in Porjus and its model, located in Älvkarleby, Sweden, have curved inlet pipes that lead the flow to the spiral casing. Nowadays, this curved pipe and its effect on the flow in the turbine is not taken into account when numerical simulations are performed at design stage. To study the impact of the inlet pipe curvature on the flow in the turbine, and to get a better overview of the flow of the whole system, measurements were made on the 1:3.1 model of the U9 turbine. Previously published measurements were taken at the inlet of the spiral casing and just before the guide vanes, using the laser Doppler anemometry (LDA) technique. In the draft tube, a number of velocity profiles were measured using the LDA techniques. The present work extends the experimental investigation with a horizontal section at the inlet of the draft tube. The experimental results are used to specify the inlet boundary condition for the numerical simulations in the draft tube, and to validate the computational results in both the spiral casing and the draft tube. The numerical simulations were realized using the standard k-e model and a block-structured hexahedral wall function mesh.

Size-selective sampling performance of six low-volume “total” suspended particulate (TSP) inlets

EPA Science Inventory

Several low-volume inlets (flow rates ≤ 16.7 liters per minute (Lpm)) are commercially available as components of low-cost, portable ambient particulate matter samplers. Because the inlets themselves do not contain internal fractionators, they are often assumed to representati...

Numerical simulation of supersonic inlets using a three-dimensional viscous flow analysis

NASA Technical Reports Server (NTRS)

Anderson, B. H.; Towne, C. E.

1980-01-01

A three dimensional fully viscous computer analysis was evaluated to determine its usefulness in the design of supersonic inlets. This procedure takes advantage of physical approximations to limit the high computer time and storage associated with complete Navier-Stokes solutions. Computed results are presented for a Mach 3.0 supersonic inlet with bleed and a Mach 7.4 hypersonic inlet. Good agreement was obtained between theory and data for both inlets. Results of a mesh sensitivity study are also shown.

Prediction of Laminar and Turbulent Boundary Layer Flow Separation in V/STOL Engine Inlets

NASA Technical Reports Server (NTRS)

Chou, D. C.; Luidens, R. W.; Stockman, N. O.

1977-01-01

A description is presented of the development of the boundary layer on the lip and diffuser surface of a subsonic inlet at arbitrary operating conditions of mass flow rate, free stream velocity and incidence angle. Both laminar separation on the lip and turbulent separation in the diffuser are discussed. The agreement of the theoretical results with model experimental data illustrates the capability of the theory to predict separation. The effects of throat Mach number, inlet size, and surface roughness on boundary layer development and separation are illustrated.

Improved computer programs for calculating potential flow in propulsion system inlets

NASA Technical Reports Server (NTRS)

Stockman, N. O.; Farrell, C. A., Jr.

1977-01-01

Computer programs to calculate the incompressible potential flow corrected for compressibility in axisymmetric inlets at arbitrary operating conditions are presented. Included are a statement of the problem to be solved, a description of each of the programs and sufficient documentation, including a test case, to enable a user to run the programs.

STUDY PROGRAM FOR TURBO-COOLER FOR PRODUCING ENGINE COOLING AIR.

DTIC Science & Technology

VANES , STAGNATION POINT, DECELERATION, ACCELERATION, SUPERSONIC DIFFUSERS, TURBINE BLADES , EVAPOTRANSPIRATION, LIQUID COOLED, HEAT TRANSFER, GAS BEARINGS, SEALS...HYPERSONIC AIRCRAFT , COOLING + VENTILATING EQUIPMENT), (*GAS TURBINES , COOLING + VENTILATING EQUIPMENT), HYPERSONIC FLOW, AIR COOLED, AIRCRAFT ... ENGINES , FEASIBILITY STUDIES, PRESSURE, SUPERSONIC CHARACTERISTICS, DESIGN, HEAT EXCHANGERS, COOLING (U) AXIAL FLOW TURBINES , DUCT INLETS, INLET GUIDE

40 CFR 60.54 - Test methods and procedures.

Code of Federal Regulations, 2011 CFR

2011-07-01

... sample CO2 concentrations at all traverse points. (ii) If sampling is conducted after a wet scrubber, an... volumetric flow rates at the inlet and outlet of the wet scrubber and the inlet CO2 concentration may be used... concentration measured before the scrubber, percent dry basis. Qdi=volumetric flow rate of effluent gas before...

Internal aerodynamics of a generic three-dimensional scramjet inlet at Mach 10

NASA Technical Reports Server (NTRS)

Holland, Scott D.

1995-01-01

A combined computational and experimental parametric study of the internal aerodynamics of a generic three-dimensional sidewall compression scramjet inlet configuration at Mach 10 has been performed. The study was designed to demonstrate the utility of computational fluid dynamics as a design tool in hypersonic inlet flow fields, to provide a detailed account of the nature and structure of the internal flow interactions, and to provide a comprehensive surface property and flow field database to determine the effects of contraction ratio, cowl position, and Reynolds number on the performance of a hypersonic scramjet inlet configuration. The work proceeded in several phases: the initial inviscid assessment of the internal shock structure, the preliminary computational parametric study, the coupling of the optimized configuration with the physical limitations of the facility, the wind tunnel blockage assessment, and the computational and experimental parametric study of the final configuration. Good agreement between computation and experimentation was observed in the magnitude and location of the interactions, particularly for weakly interacting flow fields. Large-scale forward separations resulted when the interaction strength was increased by increasing the contraction ratio or decreasing the Reynolds number.

Preliminary Results of an Altitude-Wind-Tunnel Investigation of an Axial-Flow Gas Turbine-Propeller Engine. 4; Compressor and Turbine Performance Characteristics

NASA Technical Reports Server (NTRS)

Wallner, Lewis E.; Saari, Martin J.

1948-01-01

As part of an investigation of the performance and operational characteristics of the axial-flow gas turbine-propeller engine, conducted in the Cleveland altitude wind tunnel, the performance characteristics of the compressor and the turbine were obtained. The data presented were obtained at a compressor-inlet ram-pressure ratio of 1.00 for altitudes from 5000 to 35,000 feet, engine speeds from 8000 to 13,000 rpm, and turbine-inlet temperatures from 1400 to 2100 R. The highest compressor pressure ratio obtained was 6.15 at a corrected air flow of 23.7 pounds per second and a corrected turbine-inlet temperature of 2475 R. Peak adiabatic compressor efficiencies of about 77 percent were obtained near the value of corrected air flow corresponding to a corrected engine speed of 13,000 rpm. This maximum efficiency may be somewhat low, however, because of dirt accumulations on the compressor blades. A maximum adiabatic turbine efficiency of 81.5 percent was obtained at rated engine speed for all altitudes and turbine-inlet temperatures investigated.

Characterization of flow in a scroll duct

NASA Technical Reports Server (NTRS)

Begg, E. K.; Bennett, J. C.

1985-01-01

A quantitative, flow visualization study was made of a partially elliptic cross section, inward curving duct (scroll duct), with an axial outflow through a vaneless annular cutlet. The working fluid was water, with a Re(d) of 40,000 at the inlet to the scroll duct, this Reynolds number being representative of the conditions in an actual gas turbine scroll. Both still and high speed moving pictures of fluorescein dye injected into the flow and illuminated by an argon ion laser were used to document the flow. Strong secondary flow, similar to the secondary flow in a pipe bend, was found in the bottom half of the scroll within the first 180 degs of turning. The pressure field set up by the turning duct was strong enough to affect the inlet flow condition. At 90 degs downstream, the large scale secondary flow was found to be oscillatory in nature. The exit flow was nonuniform in the annular exit. By 270 degs downstream, the flow appeared unorganized with no distinctive secondary flow pattern. Large scale structures from the upstream core region appeared by 90 degs and continued through the duct to reenter at the inlet section.

NASCRIN - NUMERICAL ANALYSIS OF SCRAMJET INLET

NASA Technical Reports Server (NTRS)

Kumar, A.

1994-01-01

The NASCRIN program was developed for analyzing two-dimensional flow fields in supersonic combustion ramjet (scramjet) inlets. NASCRIN solves the two-dimensional Euler or Navier-Stokes equations in conservative form by an unsplit, explicit, two-step finite-difference method. A more recent explicit-implicit, two-step scheme has also been incorporated in the code for viscous flow analysis. An algebraic, two-layer eddy-viscosity model is used for the turbulent flow calculations. NASCRIN can analyze both inviscid and viscous flows with no struts, one strut, or multiple struts embedded in the flow field. NASCRIN can be used in a quasi-three-dimensional sense for some scramjet inlets under certain simplifying assumptions. Although developed for supersonic internal flow, NASCRIN may be adapted to a variety of other flow problems. In particular, it should be readily adaptable to subsonic inflow with supersonic outflow, supersonic inflow with subsonic outflow, or fully subsonic flow. The NASCRIN program is available for batch execution on the CDC CYBER 203. The vectorized FORTRAN version was developed in 1983. NASCRIN has a central memory requirement of approximately 300K words for a grid size of about 3,000 points.

Nonlinear optimal control policies for buoyancy-driven flows in the built environment

NASA Astrophysics Data System (ADS)

Nabi, Saleh; Grover, Piyush; Caulfield, Colm

2017-11-01

We consider optimal control of turbulent buoyancy-driven flows in the built environment, focusing on a model test case of displacement ventilation with a time-varying heat source. The flow is modeled using the unsteady Reynolds-averaged equations (URANS). To understand the stratification dynamics better, we derive a low-order partial-mixing ODE model extending the buoyancy-driven emptying filling box problem to the case of where both the heat source and the (controlled) inlet flow are time-varying. In the limit of a single step-change in the heat source strength, our model is consistent with that of Bower et al.. Our model considers the dynamics of both `filling' and `intruding' added layers due to a time-varying source and inlet flow. A nonlinear direct-adjoint-looping optimal control formulation yields time-varying values of temperature and velocity of the inlet flow that lead to `optimal' time-averaged temperature relative to appropriate objective functionals in a region of interest.

ON THE MODIFICATION OF THE LOW FLOW-RATE PM10 DICHOTOMOUS SAMPLER INLET

EPA Science Inventory

A popular flat-topped inlet used for the collection of atmospheric particulate matter was modified to reduce water intrusion during rain and snow events. Simple alterations in the intake region of this inlet were made, including a larger drain hole, a one piece top plate, and ...

Aerosol and trace gas flux measurements from a mobile car platform on the highway

NASA Astrophysics Data System (ADS)

Gordon, M.; Miller, S. J.; Staebler, R. M.; Taylor, P.

2016-12-01

Mobile flux measurements of aerosols and trace gases at the surface can provide valuable information about the vertical transport of these compounds from near-surface sources. These measurements can be complimentary to stationary tower measurements or elevated mobile measurements from aircraft and unmanned aerial systems (UAS). In July, 2016 a mobile platform (Toyota Highlander), outfitted with a sonic anemometer (ATI), an open path CO2/H2O analyzer (Licor), and an ultrafine particle sizer (DMT), was driven on highways as part of a chasing study to investigate vehicle-induced turbulence and mixing. The open path analyzer and particle sizer inlet were co-located with the anemometer in order to investigate the feasibility of making flux measurements of heat, momentum, water vapour, CO2, and sub-micron aerosols on the highway. These highway flux measurements are compared to stationary platform measurements made upwind and downwind of the highway. Statistical and spectral analyses are used to demonstrate the validity of the mobile measurements. Uncertainties due to flow distortion around the vehicle, under-sampling, and heterogeneity of the vertical temperature and concentrations are investigated and discussed.

Mean flowfields in axisymmetric combustor geometries with swirl

NASA Astrophysics Data System (ADS)

Rhode, D. L.; Lilley, D. G.; McLaughlin, D. K.

1982-01-01

Six flowfield configurations are investigated with sidewall angles of 90 and 45 deg, and swirl vane angles of 0, 45, and 70 deg. It is found that central recirculation zones occur for the swirling flow cases investigated, which extend from the inlet to x/D = 1.7, where x is the axial polar coordinate, and D is the test section diameter. Five-hole pitot probe pressure measurements are used to determine time-mean velocities, and corresponding flow situations are predicted and compared to results of experimental data. Excellent agreement is found for the nonswirling flow, although poor agreement is found for swirling flow cases, especially near the inlet. The discrepancy is attributed to the lack of realism in the turbulence model, and/or to inaccurate specification of time-mean velocity and turbulence energy distributions at the inlet.

Low-speed performance of an axisymmetric, mixed-compression, supersonic inlet with auxiliary inlets

NASA Technical Reports Server (NTRS)

Trefny, C. J.; Wasserbauer, J. W.

1986-01-01

A test program was conducted to determine the aerodynamic performance and acoustic characteristics associated with the low-speed operation of a supersonic, axisymmetric, mixed-compression inlet with auxiliary inlets. Blow-in-auxiliary doors were installed on the NASA Ames P inlet. One door per quadrant was located on the cowl in the subsonic diffuser selection of the inlet. Auxiliary inlets with areas of 20 and 40 percent of the inlet capture area were tested statically and at free-stream Mach numbers of 0.1 and 0.2. The effects of boundary layer bleed inflow were investigated. A JT8D fan simulator driven by compressed air was used to pump inlet flow and to provide a characteristic noise signature. Baseline data were obtained at static free-stream conditions with the sharp P-inlet cowl lip replaced by a blunt lip. Auxiliary inlets increased overall total pressure recovery of the order of 10 percent.

Off-Design Performance of a Streamline-Traced, External-Compression Supersonic Inlet

NASA Technical Reports Server (NTRS)

Slater, John W.

2017-01-01

A computational study was performed to explore the aerodynamic performance of a streamline-traced, external-compression inlet designed for Mach 1.664 at off-design conditions of freestream Mach number, angle-of-attack, and angle-of-sideslip. Serious degradation of the inlet performance occurred for negative angles-of-attack and angles-of-sideslip greater than 3 degrees. At low subsonic speeds, the swept leading edges of the inlet created a pair of vortices that propagated to the engine face. Increasing the bluntness of the cowl lip showed no real improvement in the inlet performance at the low speeds, but did improve the inlet performance at the design conditions. Reducing the inlet flow rate improved the inlet performance, but at the likely expense of reduced thrust of the propulsion system. Deforming the cowl lip for low-speed operation of the inlet increased the inlet capture area and improved the inlet performance.

A Lagrangian Approach for Calculating Microsphere Deposition in a One-Dimensional Lung-Airway Model.

PubMed

Vaish, Mayank; Kleinstreuer, Clement

2015-09-01

Using the open-source software openfoam as the solver, a novel approach to calculate microsphere transport and deposition in a 1D human lung-equivalent trumpet model (TM) is presented. Specifically, for particle deposition in a nonlinear trumpetlike configuration a new radial force has been developed which, along with the regular drag force, generates particle trajectories toward the wall. The new semi-empirical force is a function of any given inlet volumetric flow rate, micron-particle diameter, and lung volume. Particle-deposition fractions (DFs) in the size range from 2 μm to 10 μm are in agreement with experimental datasets for different laminar and turbulent inhalation flow rates as well as total volumes. Typical run times on a single processor workstation to obtain actual total deposition results at comparable accuracy are 200 times less than that for an idealized whole-lung geometry (i.e., a 3D-1D model with airways up to 23rd generation in single-path only).

A Combined CFD/Characteristic Method for Prediction and Design of Hypersonic Inlet with Nose Bluntness

NASA Astrophysics Data System (ADS)

Gao, Wenzhi; Li, Zhufei; Yang, Jiming

Leading edge bluntness is widely used in hypersonic inlet design for thermal protection[1]. Detailed research of leading edge bluntness on hypersonic inlet has been concentrated on shock shape correlation[2], boundary layer flow[3], inlet performance[4], etc. It is well known that blunted noses cause detached bow shocks which generate subsonic regions around the noses and entropy layers in the flowfield.

Thermoelectric Generation Using Counter-Flows of Ideal Fluids

NASA Astrophysics Data System (ADS)

Meng, Xiangning; Lu, Baiyi; Zhu, Miaoyong; Suzuki, Ryosuke O.

2017-08-01

Thermoelectric (TE) performance of a three-dimensional (3-D) TE module is examined by exposing it between a pair of counter-flows of ideal fluids. The ideal fluids are thermal sources of TE module flow in the opposite direction at the same flow rate and generate temperature differences on the hot and cold surfaces due to their different temperatures at the channel inlet. TE performance caused by different inlet temperatures of thermal fluids are numerically analyzed by using the finite-volume method on 3-D meshed physical models and then compared with those using a constant boundary temperature. The results show that voltage and current of the TE module increase gradually from a beginning moment to a steady flow and reach a stable value. The stable values increase with inlet temperature of the hot fluid when the inlet temperature of cold fluid is fixed. However, the time to get to the stable values is almost consistent for all the temperature differences. Moreover, the trend of TE performance using a fluid flow boundary is similar to that of using a constant boundary temperature. Furthermore, 3-D contours of fluid pressure, temperature, enthalpy, electromotive force, current density and heat flux are exhibited in order to clarify the influence of counter-flows of ideal fluids on TE generation. The current density and heat flux homogeneously distribute on an entire TE module, thus indicating that the counter-flows of thermal fluids have high potential to bring about fine performance for TE modules.

The change of the inlet geometry of a centrifugal compressor stage and its influence on the compressor performance

NASA Astrophysics Data System (ADS)

Wang, Leilei; Yang, Ce; Zhao, Ben; Lao, Dazhong; Ma, Chaochen; Li, Du

2013-06-01

The impact on the compressor performance is important for designing the inlet pipe of the centrifugal compressor of a vehicle turbocharger with different inlet pipes. First, an experiment was performed to determine the compressor performance from three cases: a straight inlet pipe, a long bent inlet pipe and a short bent inlet pipe. Next, dynamic sensors were installed in key positions to collect the sign of the unsteady pressure of the centrifugal compressor. Combined with the results of numerical simulations, the total pressure distortion in the pipes, the pressure distributions on the blades and the pressure variability in the diffuser are studied in detail. The results can be summarized as follows: a bent pipe results in an inlet distortion to the compressor, which leads to performance degradation, and the effect is more apparent as the mass flow rate increases. The distortion induced by the bent inlet is not only influenced by the distance between the outlet of the bent section and the leading edge of the impeller but also by the impeller rotation. The flow fields in the centrifugal impeller and the diffuser are influenced by a coupling effect produced by the upstream inlet distortion and the downstream blocking effect from the volute tongue. If the inlet geometry is changed, the distributions and the fluctuation intensities of the static pressure on the main blade surface of the centrifugal impeller and in the diffuser are changed accordingly.

Analysis of a Channeled Centerbody Supersonic Inlet for F-15B Flight Research

NASA Technical Reports Server (NTRS)

Ratnayake, Nalin A.

2010-01-01

The Propulsion Flight Test Fixture at the NASA Dryden Flight Research Center is a unique test platform available for use on the NASA F-15B airplane, tail number 836, as a modular host for a variety of aerodynamics and propulsion research. The first experiment that is to be flown on the test fixture is the Channeled Centerbody Inlet Experiment. The objectives of this project at Dryden are twofold: 1) flight evaluation of an innovative new approach to variable geometry for high-speed inlets, and 2) flight validation of channeled inlet performance prediction by complex computational fluid dynamics codes. The inlet itself is a fixed-geometry version of a mixed-compression, variable-geometry, supersonic in- let developed by TechLand Research, Inc. (North Olmsted, Ohio) to improve the efficiency of supersonic flight at off-nominal conditions. The concept utilizes variable channels in the centerbody section to vary the mass flow of the inlet, enabling efficient operation at a range of flight conditions. This study is particularly concerned with the starting characteristics of the inlet. Computational fluid dynamics studies were shown to align well with analytical predictions, showing the inlet to remain unstarted as designed at the primary test point of Mach 1.5 at an equivalent pressure altitude of 29,500 ft local conditions. Mass-flow-related concerns such as the inlet start problem, as well as inlet efficiency in terms of total pressure loss, are assessed using the flight test geometry.

Local flow measurements at the inlet spike tip of a Mach 3 supersonic cruise airplane

NASA Technical Reports Server (NTRS)

Johnson, H. J.; Montoya, E. J.

1973-01-01

The flow field at the left inlet spike tip of a YF-12A airplane was examined using at 26 deg included angle conical flow sensor to obtain measurements at free-stream Mach numbers from 1.6 to 3.0. Local flow angularity, Mach number, impact pressure, and mass flow were determined and compared with free-stream values. Local flow changes occurred at the same time as free-stream changes. The local flow usually approached the spike centerline from the upper outboard side because of spike cant and toe-in. Free-stream Mach number influenced the local flow angularity; as Mach number increased above 2.2, local angle of attack increased and local sideslip angle decreased. Local Mach number was generally 3 percent less than free-stream Mach number. Impact-pressure ratio and mass flow ratio increased as free-stream Mach number increased above 2.2, indicating a beneficial forebody compression effect. No degradation of the spike tip instrumentation was observed after more than 40 flights in the high-speed thermal environment encountered by the airplane. The sensor is rugged, simple, and sensitive to small flow changes. It can provide accurate imputs necessary to control an inlet.

Rankine cycle condenser pressure control using an energy conversion device bypass valve

DOEpatents

Ernst, Timothy C; Nelson, Christopher R; Zigan, James A

2014-04-01

The disclosure provides a waste heat recovery system and method in which pressure in a Rankine cycle (RC) system of the WHR system is regulated by diverting working fluid from entering an inlet of an energy conversion device of the RC system. In the system, an inlet of a controllable bypass valve is fluidly coupled to a working fluid path upstream of an energy conversion device of the RC system, and an outlet of the bypass valve is fluidly coupled to the working fluid path upstream of the condenser of the RC system such that working fluid passing through the bypass valve bypasses the energy conversion device and increases the pressure in a condenser. A controller determines the temperature and pressure of the working fluid and controls the bypass valve to regulate pressure in the condenser.

Development of a solenoid actuated planar valveless micropump with single and multiple inlet-outlet arrangements

NASA Astrophysics Data System (ADS)

Kumar, N.; George, D.; Sajeesh, P.; Manivannan, P. V.; Sen, A. K.

2016-07-01

We report a planar solenoid actuated valveless micropump with multiple inlet-outlet configurations. The self-priming characteristics of the multiple inlet-multiple outlet micropump are studied. The filling dynamics of the micropump chamber during start-up and the effects of fluid viscosity, voltage and frequency on the dynamics are investigated. Numerical simulations for multiple inlet-multiple outlet micropumps are carried out using fluid structure algorithm. With DI water and at 5.0 Vp-p, 20 Hz frequency, the two inlet-two outlet micropump provides a maximum flow rate of 336 μl min-1 and maximum back pressure of 441 Pa. Performance characteristics of the two inlet-two outlet micropump are studied for aqueous fluids of different viscosity. Transport of biological cell lines and diluted blood samples are demonstrated; the flow rate-frequency characteristics are studied. Viability of cells during pumping with multiple inlet multiple outlet configuration is also studied in this work, which shows 100% of cells are viable. Application of the proposed micropump for simultaneous pumping, mixing and distribution of fluids is demonstrated. The proposed integrated, standalone and portable micropump is suitable for drug delivery, lab-on-chip and micro-total-analysis applications.

An Investigation of the Effects of Nose and Lip Shapes for an Underslung Scoop Inlet at Mach Numbers from 0 to 1.9

NASA Technical Reports Server (NTRS)

Pfyl, Frank A.

1955-01-01

An experimental investigation was conducted to determine the performance characteristics an underslung nose-scoop air-induction system for a supersonic airplane. Five different nose shapes, three lip shapes, and two internal diffusers were investigated. Tests were made at Mach numbers from 0 to 1.9, angles of attack from 0 deg to approximately l5 deg, and mass-flow ratios from 0 to maximum obtainable. It was found that the underslung nose-scoop inlet was able to operate at Mach numbers from 0.6 to 1.9 over a large positive angle-of-attack range without adverse effects on the pressure recovery. Although there was no one inlet configuration that was markedly superior over the entire range of operating variables, the arrangement having a nose designed to give increased supersonic compression at low angles of attack, and a sharp lip (configuration designated N3L3) showed the most favorable performance characteristics over the supersonic Mach number range. Inlets with sizable lip radii gave satisfactory performance up to a Mach number of 1.5; however, as a result of an increase in drag, the performance of such inlets was markedly inferior to the sharp-lip configuration above Mach numbers of 1.5. Throughout the range of test Mach numbers all inlet configurations evidenced stable air-flow characteristics over the mass-flow range for normal engine operation. Analysis of the inlet performance on the basis of a propulsive thrust parameter showed that a fixed inlet area could be used for Mach numbers up to 1.5 with only a small sacrifice in performance.

Quality by Design approach to spray drying processing of crystalline nanosuspensions.

PubMed

Kumar, Sumit; Gokhale, Rajeev; Burgess, Diane J

2014-04-10

Quality by Design (QbD) principles were explored to understand spray drying process for the conversion of liquid nanosuspensions into solid nano-crystalline dry powders using indomethacin as a model drug. The effects of critical process variables: inlet temperature, flow and aspiration rates on critical quality attributes (CQAs): particle size, moisture content, percent yield and crystallinity were investigated employing a full factorial design. A central cubic design was employed to generate the response surface for particle size and percent yield. Multiple linear regression analysis and ANOVA were employed to identify and estimate the effect of critical parameters, establish their relationship with CQAs, create design space and model the spray drying process. Inlet temperature was identified as the only significant factor (p value <0.05) to affect dry powder particle size. Higher inlet temperatures caused drug surface melting and hence aggregation of the dried nano-crystalline powders. Aspiration and flow rates were identified as significant factors affecting yield (p value <0.05). Higher yields were obtained at higher aspiration and lower flow rates. All formulations had less than 3% (w/w) moisture content. Formulations dried at higher inlet temperatures had lower moisture compared to those dried at lower inlet temperatures. Published by Elsevier B.V.

Computer programs for calculating two-dimensional potential flow in and about propulsion system inlets

NASA Technical Reports Server (NTRS)

Hawk, J. D.; Stockman, N. O.; Farrell, C. A., Jr.

1978-01-01

Incompressible potential flow calculations are presented that were corrected for compressibility in two-dimensional inlets at arbitrary operating conditions. Included are a statement of the problem to be solved, a description of each of the computer programs, and sufficient documentation, including a test case, to enable a user to run the program.

Standardized performance tests of collectors of solar thermal energy: An evacuated flatplate copper collector with a serpentine flow distribution

NASA Technical Reports Server (NTRS)

Johnson, S. M.

1976-01-01

Basic test results are given for a flat plate solar collector whose performance was determined in the NASA-Lewis solar simulator. The collector was tested over ranges of inlet temperatures, fluxes and one coolant flow rate. Collector efficiency is correlated in terms of inlet temperature and flux level.

Hydrodynamics of a Multistage Wet Scrubber Incineration Conditions

ERIC Educational Resources Information Center

Said, M. M.; Manyele, S. V.; Raphael, M. L.

2012-01-01

The objective of the study was to determine the hydrodynamics of the two stage counter-current cascade wet scrubbers used during incineration of medical waste. The dependence of the hydrodynamics on two main variables was studied: Inlet air flow rate and inlet liquid flow rate. This study introduces a new wet scrubber operating features, which are…

Effect of inlet conditions for numerical modelling of the urban boundary layer

NASA Astrophysics Data System (ADS)

Gnatowska, Renata

2018-01-01

The paper presents the numerical results obtained with the use of the ANSYS FLUENT commercial code for analysing the flow structure around two rectangular inline surface-mounted bluff bodies immersed in a boundary layer. The effects of the inflow boundary layer for the accuracy of the numerical modelling of the flow field around a simple system of objects are described. The analysis was performed for two concepts. In the former case, the inlet velocity profile was defined using the power law, whereas the kinetic and dissipation energy was defined from the equations according to Richards and Hoxey [1]. In the latter case, the inlet conditions were calculated for the flow over the rough area composed of the rectangular components.

Partially turbulated trailing edge cooling passages for gas turbine nozzles

DOEpatents

Thatcher, Jonathan Carl; Burdgick, Steven Sebastian

2001-01-01

A plurality of passages are spaced one from the other along the length of a trailing edge of a nozzle vane in a gas turbine. The passages lie in communication with a cavity in the vane for flowing cooling air from the cavity through the passages through the tip of the trailing edge into the hot gas path. Each passage is partially turbulated and includes ribs in an aft portion thereof to provide enhanced cooling effects adjacent the tip of the trailing edge. The major portions of the passages are smooth bore. By this arrangement, reduced temperature gradients across the trailing edge metal are provided. Additionally, the inlets to each of the passages have a restriction whereby a reduced magnitude of compressor bleed discharge air is utilized for trailing edge cooling purposes.

Gas turbine cooling system

DOEpatents

Bancalari, Eduardo E.

2001-01-01

A gas turbine engine (10) having a closed-loop cooling circuit (39) for transferring heat from the hot turbine section (16) to the compressed air (24) produced by the compressor section (12). The closed-loop cooling system (39) includes a heat exchanger (40) disposed in the flow path of the compressed air (24) between the outlet of the compressor section (12) and the inlet of the combustor (14). A cooling fluid (50) may be driven by a pump (52) located outside of the engine casing (53) or a pump (54) mounted on the rotor shaft (17). The cooling circuit (39) may include an orifice (60) for causing the cooling fluid (50) to change from a liquid state to a gaseous state, thereby increasing the heat transfer capacity of the cooling circuit (39).

Analysis of Buzz in a Supersonic Inlet

NASA Technical Reports Server (NTRS)

Chima, Rodrick V.

2012-01-01

A dual-stream, low-boom supersonic inlet designed for use on a small, Mach 1.6 aircraft was tested experimentally in the 8- by 6-Foot Supersonic Wind Tunnel (SWT) at the NASA Glenn Research Center (GRC). The tests showed that the inlet had good recovery and stable operation over large mass flow range. The inlet went into buzz at mass flows well below that needed for engine operation, and the experiments generated a wealth of data during buzz. High frequency response pressure measurements and high-speed schlieren videos were recorded for many buzz events. The objective of the present work was to use computational fluid dynamics (CFD) to predict some of the experimental data taken during buzz, compare those predictions to the experimental data, and to use both datasets to explain the physics of the buzz cycle. The calculations were done with the Wind-US CFD code using a second-order time-accurate differencing scheme and the SST turbulence model. Computed Mach number contours were compared with schlieren images, and ensemble-averaged unsteady pressures were compared to data. The results showed that the buzz cycle consisted partly of spike buzz, an unsteady oscillation of the main shock at the spike tip while the inlet pressure dropped, and partly of choked flow while the inlet repressurized. Most of the results could be explained by theory proposed by Dailey in 1954, but did not support commonly used acoustic resonance explanations.

Combined Microfluidic-Eectric Diffused Mixing of Living Cells in Continuous Flow

NASA Astrophysics Data System (ADS)

Ming-Wen Wang,

2010-02-01

The mixing process is a crucially important stage in the operation of biological and chemical microfluidic devices. If the mixing is inadequate, reactants do not fully interact with each other, and the device may not operate properly. This paper describes a simplified microfluidic mixer (different from a chaotic mixer) which can uniformly mix a buffer solution with living cells by applying an AC electric charge. Diffusion of the living cells into the buffer solution occurs rapidly following the interface of the flow stream with the electric charge; no further agitating step is needed. To accomplish this, an asymmetric pair of electrodes was integrated at the inlets of the buffer solution and the cells fluid. When the buffer solution and the cells fluid were introduced into one flow path, they remained limited to that flow stream. When the electrodes were charged, however, the cells in a short distance were efficiently moved into the solution flow, and the original fluids were mixed. The mixing efficiency depends on the polarizability of the cells, and this in turn is governed by the dielectric properties of the cells, the medium, and the solvent. This micro device, capable of efficiently mixing living cells with a buffer solution, may potentially allow biological mixing to be done outside of hospitals, in facilities without biological analyzing instruments.

Effect of a rotating propeller on the separation angle of attack

NASA Technical Reports Server (NTRS)

Boldman, D. R.; Iek, C.; Hwang, D. P.; Larkin, M.; Schweiger, P.

1993-01-01

The present study represents an extension of an earlier wind tunnel experiment performed with the P&W 17-in. Advanced Ducted Propeller (ADP) Simulator operating at Mach 0.2. In order to study the effects of a rotating propeller on the inlet flow, data were obtained in the UTRC 10- by 15-Foot Large Subsonic Wind Tunnel with the same hardware and instrumentation, but with the propellar removed. These new tests were performed over a range of flow rates which duplicated flow rates in the powered simulator program. The flow through the inlet was provided by a remotely located vacuum source. A comparison of the results of this flow-through study with the previous data from the powered simulator indicated that in the conventional inlet the propeller produced an increase in the separation angle of attack between 4.0 deg at a specific flow of 22.4 lb/sec-sq ft to 2.7 deg at a higher specific flow of 33.8 lb/sec-sq ft. A similar effect on separation angle of attack was obtained by using stationary blockage rather than a propeller.

76 FR 9449 - National Emission Standards for Hazardous Air Pollutants: Gold Mine Ore Processing and Production...

Federal Register 2010, 2011, 2012, 2013, 2014

2011-02-17

... chemistry, scrubber pressure drop, and scrubber inlet gas temperature hourly. The final rule does not... pressure) and inlet gas temperature to be based on the minimum flow rate (or line pressure) or maximum inlet gas temperature established during the initial performance test. It also includes two additional...

Spiral blood flows in an idealized 180-degree curved artery model

NASA Astrophysics Data System (ADS)

Bulusu, Kartik V.; Kulkarni, Varun; Plesniak, Michael W.

2017-11-01

Understanding of cardiovascular flows has been greatly advanced by the Magnetic Resonance Velocimetry (MRV) technique and its potential for three-dimensional velocity encoding in regions of anatomic interest. The MRV experiments were performed on a 180-degree curved artery model using a Newtonian blood analog fluid at the Richard M. Lucas Center at Stanford University employing a 3 Tesla General Electric (Discovery 750 MRI system) whole body scanner with an eight-channel cardiac coil. Analysis in two regions of the model-artery was performed for flow with Womersley number=4.2. In the entrance region (or straight-inlet pipe) the unsteady pressure drop per unit length, in-plane vorticity and wall shear stress for the pulsatile, carotid artery-based flow rate waveform were calculated. Along the 180-degree curved pipe (curvature ratio =1/7) the near-wall vorticity and the stretching of the particle paths in the vorticity field are visualized. The resultant flow behavior in the idealized curved artery model is associated with parameters such as Dean number and Womersley number. Additionally, using length scales corresponding to the axial and secondary flow we attempt to understand the mechanisms leading to the formation of various structures observed during the pulsatile flow cycle. Supported by GW Center for Biomimetics and Bioinspired Engineering (COBRE), MRV measurements in collaboration with Prof. John K. Eaton and, Dr. Chris Elkins at Stanford University.

High efficiency virtual impactor

DOEpatents

Loo, B.W.

1980-03-27

Environmental monitoring of atmospheric air is facilitated by a single stage virtual impactor for separating an inlet flow (Q/sub 0/) having particulate contaminants into a coarse particle flow (Q/sub 1/) and a fine particle flow (Q/sub 2/) to enable collection of such particles on different filters for separate analysis. An inlet particle acceleration nozzle and coarse particle collection probe member having a virtual impaction opening are aligned along a single axis and spaced apart to define a flow separation region at which the fine particle flow (Q/sub 2/) is drawn radially outward into a chamber while the coarse particle flow (Q/sub 1/) enters the virtual impaction opening.

Nearly Interactive Parabolized Navier-Stokes Solver for High Speed Forebody and Inlet Flows

NASA Technical Reports Server (NTRS)

Benson, Thomas J.; Liou, May-Fun; Jones, William H.; Trefny, Charles J.

2009-01-01

A system of computer programs is being developed for the preliminary design of high speed inlets and forebodies. The system comprises four functions: geometry definition, flow grid generation, flow solver, and graphics post-processor. The system runs on a dedicated personal computer using the Windows operating system and is controlled by graphical user interfaces written in MATLAB (The Mathworks, Inc.). The flow solver uses the Parabolized Navier-Stokes equations to compute millions of mesh points in several minutes. Sample two-dimensional and three-dimensional calculations are demonstrated in the paper.

A Comparison of Hybrid Reynolds Averaged Navier Stokes/Large Eddy Simulation (RANS/LES) and Unsteady RANS Predictions of Separated Flow for a Variable Speed Power Turbine Blade Operating with Low Inlet Turbulence Levels

DTIC Science & Technology

2017-10-01

Facility is a large-scale cascade that allows detailed flow field surveys and blade surface measurements.10–12 The facility has a continuous run ...structured grids at 2 flow conditions, cruise and takeoff, of the VSPT blade . Computations were run in parallel on a Department of Defense...RANS/LES) and Unsteady RANS Predictions of Separated Flow for a Variable-Speed Power- Turbine Blade Operating with Low Inlet Turbulence Levels

A numerical simulation of the NFAC (National Full-scale Aerodynamics Complex) open-return wind tunnel inlet flow

NASA Technical Reports Server (NTRS)

Kaul, U. K.; Ross, J. C.; Jacocks, J. L.

1985-01-01

The flow into an open return wind tunnel inlet was simulated using Euler equations. An explicit predictor-corrector method was employed to solve the system. The calculation is time-accurate and was performed to achieve a steady-state solution. The predictions are in reasonable agreement with the experimental data. Wall pressures are accurately predicted except in a region of recirculating flow. Flow-field surveys agree qualitatively with laser velocimeter measurements. The method can be used in the design process for open return wind tunnels.

Inlet Diameter and Flow Volume Effects on Separation and Energy Efficiency of Hydrocyclones

NASA Astrophysics Data System (ADS)

Erikli, Ş.; Olcay, A. B.

2015-08-01

This study investigates hydrocyclone performance of an oil injected screw compressor. Especially, the oil separation efficiency of a screw compressor plays a significant role for air quality and non-stop working hour of compressors has become an important issue when the efficiency in energy is considered. In this study, two separation efficiency parameters were selected to be hydrocyclone inlet diameter and flow volume height between oil reservoir surface and top of the hydrocyclone. Nine different cases were studied in which cyclone inlet diameter and flow volume height between oil reservoir surface and top were investigated in regards to separation and energy performance aspects and the effect of the parameters on the general performance appears to be causing powerful influence. Flow inside the hydrocyclone geometry was modelled by Reynolds Stress Model (RSM) and hydro particles were tracked by Discrete Phase Model (DPM). Besides, particle break up was modelled by the Taylor Analogy Breakup (TAB) model. The reversed vortex generation was observed at different planes. The upper limit of the inlet diameter of the cyclone yields the centrifugal force on particles to decrease while the flow becomes slower; and the larger diameter implies slower flow. On the contrary, the lower limit is increment in speed causes breakup problems that the particle diameters become smaller; consequently, it is harder to separate them from gas.

Quiet Clean Short-haul Experimental Engine (QCSEE). Aerodynamic and aeromechanical performance of a 50.8 cm (20 inch) diameter 1.34 PR variable pitch fan with core flow

NASA Technical Reports Server (NTRS)

Giffin, R. G.; Mcfalls, R. A.; Beacher, B. F.

1977-01-01

The fan aerodynamic and aeromechanical performance tests of the quiet clean short haul experimental engine under the wing fan and inlet with a simulated core flow are described. Overall forward mode fan performance is presented at each rotor pitch angle setting with conventional flow pressure ratio efficiency fan maps, distinguishing the performance characteristics of the fan bypass and fan core regions. Effects of off design bypass ratio, hybrid inlet geometry, and tip radial inlet distortion on fan performance are determined. The nonaxisymmetric bypass OGV and pylon configuration is assessed relative to both total pressure loss and induced circumferential flow distortion. Reverse mode performance, obtained by resetting the rotor blades through both the stall pitch and flat pitch directions, is discussed in terms of the conventional flow pressure ratio relationship and its implications upon achievable reverse thrust. Core performance in reverse mode operation is presented in terms of overall recovery levels and radial profiles existing at the simulated core inlet plane. Observations of the starting phenomena associated with the initiation of stable rotor flow during acceleration in the reverse mode are briefly discussed. Aeromechanical response characteristics of the fan blades are presented as a separate appendix, along with a description of the vehicle instrumentation and method of data reduction.

Additional testing of the inlets designed for a tandem fan V/STOL nacelle

NASA Technical Reports Server (NTRS)

Ybarra, A. H.

1981-01-01

The wind tunnel testing of a scale model of a tandem fan nacelle designed for a type (subsonic cruise) V/STOL aircraft configuration is discussed. The performance for the isolated front inlet and for the combined front and aft inlets is reported. Model variables include front and aft inlets with aft inlet variations of short and long aft inlet cowls, with a shaft simulator and diffuser vortex generators, cowl lip fillets, and nacelle strakes. Inlet pressure recovery, distortion, and inlet angle-to-attack separation limits were evaluated at tunnel velocity from 0 to 240 knots, angles-of-attack from -10 to +40 degrees and inlet flow rates corresponding to throat Mach number from 0.0 to 0.6. Combined nacelle pitch and yaw runs up to 30 deg. were also made.

Adjustable flow rate controller for polymer solutions

DOEpatents

Jackson, Kenneth M.

1981-01-01

An adjustable device for controlling the flow rate of polymer solutions which results in only little shearing of the polymer molecules, said device comprising an inlet manifold, an outlet manifold, a plurality of tubes capable of providing communication between said inlet and outlet manifolds, said tubes each having an internal diameter that is smaller than that of the inlet manifold and large enough to insure that viscosity of the polymer solution passing through each said tube will not be reduced more than about 25 percent, and a valve associated with each tube, said valve being capable of opening or closing communication in that tube between the inlet and outlet manifolds, each said valve when fully open having a diameter that is substantially at least as great as that of the tube with which it is associated.

Indicator providing continuous indication of the presence of a specific pollutant in air

NASA Technical Reports Server (NTRS)

Miller, C. G.; Bartera, R. E. (Inventor)

1976-01-01

A continuous HCl in-air indicator was developed which consists of a tube-like element with an inlet end through which a continuous stream of air containing HCl enters. The air flows downstream from the inlet end and exits the element's outlet end. Positioned between the element's inlet and outlet ends are first and second spaced apart photoelectric units, which are preferably positioned adjacent the inlet and outlet ends, respectively. Ammonia gas is injected into the air, flowing through the element, at a position between the two photoelectric units. The ammonia gas reacts with the HCl in the air to form ammonium chloride particles. The difference between the outputs of the two photoelectric units is an indication of the amount of HCl in the air stream.

A Combined Experimental/Computational Investigation of a Rocket Based Combined Cycle Inlet

NASA Technical Reports Server (NTRS)

Smart, Michael K.; Trexler, Carl A.; Goldman, Allen L.

2001-01-01

A rocket based combined cycle inlet geometry has undergone wind tunnel testing and computational analysis with Mach 4 flow at the inlet face. Performance parameters obtained from the wind tunnel tests were the mass capture, the maximum back-pressure, and the self-starting characteristics of the inlet. The CFD analysis supplied a confirmation of the mass capture, the inlet efficiency and the details of the flowfield structure. Physical parameters varied during the test program were cowl geometry, cowl position, body-side bleed magnitude and ingested boundary layer thickness. An optimum configuration was determined for the inlet as a result of this work.

Radial flow nuclear thermal rocket (RFNTR)

DOEpatents

Leyse, Carl F.

1995-11-07

A radial flow nuclear thermal rocket fuel assembly includes a substantially conical fuel element having an inlet side and an outlet side. An annular channel is disposed in the element for receiving a nuclear propellant, and a second, conical, channel is disposed in the element for discharging the propellant. The first channel is located radially outward from the second channel, and separated from the second channel by an annular fuel bed volume. This fuel bed volume can include a packed bed of loose fuel beads confined by a cold porous inlet frit and a hot porous exit frit. The loose fuel beads include ZrC coated ZrC-UC beads. In this manner, nuclear propellant enters the fuel assembly axially into the first channel at the inlet side of the element, flows axially across the fuel bed volume, and is discharged from the assembly by flowing radially outward from the second channel at the outlet side of the element.

Radial flow nuclear thermal rocket (RFNTR)

DOEpatents

Leyse, Carl F.

1995-01-01

A radial flow nuclear thermal rocket fuel assembly includes a substantially conical fuel element having an inlet side and an outlet side. An annular channel is disposed in the element for receiving a nuclear propellant, and a second, conical, channel is disposed in the element for discharging the propellant. The first channel is located radially outward from the second channel, and separated from the second channel by an annular fuel bed volume. This fuel bed volume can include a packed bed of loose fuel beads confined by a cold porous inlet frit and a hot porous exit frit. The loose fuel beads include ZrC coated ZrC-UC beads. In this manner, nuclear propellant enters the fuel assembly axially into the first channel at the inlet side of the element, flows axially across the fuel bed volume, and is discharged from the assembly by flowing radially outward from the second channel at the outlet side of the element.

Design and Numerical Simulation of Radial Inflow Turbine Volute

NASA Astrophysics Data System (ADS)

Shah, Samip P.; Channiwala, S. A.; Kulshreshtha, D. B.; Chaudhari, Gaurang

2014-12-01

The volute of a radial inflow turbine has to be designed to ensure that the desired rotor inlet conditions like absolute Mach number, flow angle etc. are attained. For the reasonable performance of vaneless volute turbine care has to be taken for reduction in losses at an appropriate flow angle at the rotor inlet, in the direction of volute, whose function is to convert gas energy into kinetic energy and direct the flow towards the rotor inlet at an appropriate flow angle with reduced losses. In literature it was found that the incompressible approaches failed to provide free vortex and uniform flow at rotor inlet for compressible flow regimes. So, this paper describes a non-dimensional design procedure for a vaneless turbine volute for compressible flow regime and investigates design parameters, such as the distribution of area ratio and radius ratio as a function of azimuth angle. The nondimensional design is converted in dimensional form for three different volute cross sections. A commercial computational fluid dynamics code is used to develop numerical models of three different volute cross sections. From the numerical models, losses generation in the different volutes are identified and compared. The maximum pressure loss coefficient for Trapezoidal cross section is 0.1075, for Bezier-trapezoidal cross section is 0.0677 and for circular cross section is 0.0438 near tongue region, which suggested that the circular cross section will give a better efficiency than other types of volute cross sections.

Inlet design for high-speed propfans

NASA Technical Reports Server (NTRS)

Little, B. H., Jr.; Hinson, B. L.

1982-01-01

A two-part study was performed to design inlets for high-speed propfan installation. The first part was a parametric study to select promising inlet concepts. A wide range of inlet geometries was examined and evaluated - primarily on the basis of cruise thrust and fuel burn performance. Two inlet concepts were than chosen for more detailed design studies - one apropriate to offset engine/gearbox arrangements and the other to in-line arrangements. In the second part of this study, inlet design points were chosen to optimize the net installed thrust, and detailed design of the two inlet configurations was performed. An analytical methodology was developed to account for propfan slipstream effects, transonic flow efects, and three-dimensional geometry effects. Using this methodology, low drag cowls were designed for the two inlets.

Investigation of the column performance of cadmium(II) biosorption by Cladophora crispata flocs in a packed bed

DOE Office of Scientific and Technical Information (OSTI.GOV)

Aksu, Z.; Kutsal, T.; Caglar, A.

1998-03-01

In this study the biosorption of cadmium(II) ions to dried flocs of Cladophora crispata, a kind of green algae, was investigated in a packed bed column. The cadmium(II) removal performance of the column was investigated as a function of the cadmium(II)-bearing solution flow rate and the inlet cadmium(II) concentration. Removal and total removal percentages of cadmium(II) related to flow volume were determined by evaluating the breakthrough curves obtained at three different flow rates for two different constant inlet concentrations. At the lowest flow rate the effect of inlet cadmium(II) concentration on the column capacity was also investigated. Data confirmed thatmore » early saturation and lower cadmium(II) removals were observed at higher flow rates and at higher cadmium(II) concentrations. Column experiments also showed that maximum specific cadmium(II) uptake values of C. crispata flocs were as high as those of other biomass sorbents.« less

Standardized performance tests of collectors of solar thermal energy-a flat-plate collector with a single-tube serpentine flow distribution

NASA Technical Reports Server (NTRS)

Johnson, S.

1976-01-01

This preliminary data report gives basic test results of a flat-plate solar collector whose performance was determined in the NASA-Lewis solar simulator. The collector was tested over ranges of inlet temperatures, fluxes and coolant flow rates. Collector efficienty is correlated in terms of inlet temperature and flux level.

Combined buoyancy and flow direction effects on saturated boiling critical heat flux in liquid nitrogen

NASA Technical Reports Server (NTRS)

Papell, S. S.

1972-01-01

Buoyancy effects on the critical heat flux and general data trends for a liquid nitrogen internal flow system were determined by comparison of upflow and downflow data under identical test conditions. The test section had a 1.28 cm diameter flow passage and a 30.5 cm heated length which was subjected to uniform heat fluxes through resistance heating. Test conditions covered a range of pressures from 3.4 to 10.2 atm, inlet velocities from 0.23 to 3.51 m/sec, with the liquid nitrogen temperature at saturated inlet conditions. Data comparisons showed that the critical heat flux for downflow could be up to 36 percent lower than for upflow. A nonmonotonic relationship between the critical heat flux and velocity was determined for upflow but not for downflow. A limiting inlet velocity of 4.12 m/sec was determined to be the minimum velocity required to completely suppress the influence of buoyancy on the critical heat flux for this saturated inlet flow system. A correlation of this limiting fluid velocity is presented that was developed from previously published subcooled liquid nitrogen data and the saturated data of this investigation.

The aerodynamic characteristics of vortex ingestion for the F/A-18 inlet duct

NASA Technical Reports Server (NTRS)

Anderson, Bernhard H.

1991-01-01

A Reduced Navier-Stokes (RNS) solution technique was successfully combined with the concept of partitioned geometry and mesh generation to form a very efficient 3D RNS code aimed at the analysis-design engineering environment. Partitioned geometry and mesh generation is a pre-processor to augment existing geometry and grid generation programs which allows the solver to (1) recluster an existing gridlife mesh lattice, and (2) perturb an existing gridfile definition to alter the cross-sectional shape and inlet duct centerline distribution without returning to the external geometry and grid generator. The present results provide a quantitative validation of the initial value space marching 3D RNS procedure and demonstrates accurate predictions of the engine face flow field, with a separation present in the inlet duct as well as when vortex generators are installed to supress flow separation. The present results also demonstrate the ability of the 3D RNS procedure to analyze the flow physics associated with vortex ingestion in general geometry ducts such as the F/A-18 inlet. At the conditions investigated, these interactions are basically inviscid like, i.e., the dominant aerodynamic characteristics have their origin in inviscid flow theory.

Axisymmetric Calculations of a Low-Boom Inlet in a Supersonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Chima, Rodrick V.; Hirt, Stefanie M.; Reger, Robert

2011-01-01

This paper describes axisymmetric CFD predictions made of a supersonic low-boom inlet with a facility diffuser, cold pipe, and mass flow plug within wind tunnel walls, and compares the CFD calculations with the experimental data. The inlet was designed for use on a small supersonic aircraft that would cruise at Mach 1.6, with a Mach number over the wing of 1.7. The inlet was tested in the 8-ft by 6-ft Supersonic Wind Tunnel at NASA Glenn Research Center in the fall of 2010 to demonstrate the performance and stability of a practical flight design that included a novel bypass duct. The inlet design is discussed here briefly. Prior to the test, CFD calculations were made to predict the performance of the inlet and its associated wind tunnel hardware, and to estimate flow areas needed to throttle the inlet. The calculations were done with the Wind-US CFD code and are described in detail. After the test, comparisons were made between computed and measured shock patterns, total pressure recoveries, and centerline pressures. The results showed that the dual-stream inlet had excellent performance, with capture ratios near one, a peak core total pressure recovery of 96 percent, and a large stable operating range. Predicted core recovery agreed well with the experiment but predicted bypass recovery and maximum capture ratio were high. Calculations of offdesign performance of the inlet along a flight profile agreed well with measurements and previous calculations.

Determination of shelf heat transfer coefficients along the shelf flow path of a freeze dryer using the shelf fluid temperature perturbation approach.

PubMed

Kuu, Wei Y; Nail, Steven L; Hardwick, Lisa M

2007-01-01

The spatial distribution of local shelf heat transfer coefficients, Ks, was determined by mapping the transient temperature response of the shelf surface along the serpentine internal channels of the shelf while the temperature of the heat transfer fluid was ramped from -40 degrees to 40 degrees C. The solution of a first-order non-steady-state differential equation resulted in a predicted shelf surface temperature as a function of the shelf fluid temperature at any point along the flow path. During the study, the shelf surfaces were maintained under a thermally insulated condition so that the heat transfers by gas conduction and radiation were negligible. To minimize heat conduction by gas, the chamber was evacuated to a low pressure, such as 100 mTorr. To minimize heat transfers between shelves, shelves were moved close together, with a gap of approximately 3 mm between any two shelves, because the shelf surface temperatures at corresponding vertical locations of two shelves are virtually equal. In addition, this also provides a shielding from radiation heat transfer from shelf to walls. Local heat transfer coefficients at the probed locations h(x) ( approximately Ks) were calculated by fitting the experimental shelf temperature response to the theoretical value. While the resulting values of K(s) are in general agreement with previously reported values, the values of Ks close to the inlet are significantly higher than those of other locations of the shelf channel. This observation is most likely attributed to the variation of the flow pattern of heat transfer fluid within the channels.

A semi-analytical model for the acoustic impedance of finite length circular holes with mean flow

NASA Astrophysics Data System (ADS)

Yang, Dong; Morgans, Aimee S.

2016-12-01

The acoustic response of a circular hole with mean flow passing through it is highly relevant to Helmholtz resonators, fuel injectors, perforated plates, screens, liners and many other engineering applications. A widely used analytical model [M.S. Howe. "Onthe theory of unsteady high Reynolds number flow through a circular aperture", Proc. of the Royal Soc. A. 366, 1725 (1979), 205-223] which assumes an infinitesimally short hole was recently shown to be insufficient for predicting the impedance of holes with a finite length. In the present work, an analytical model based on Green's function method is developed to take the hole length into consideration for "short" holes. The importance of capturing the modified vortex noise accurately is shown. The vortices shed at the hole inlet edge are convected to the hole outlet and further downstream to form a vortex sheet. This couples with the acoustic waves and this coupling has the potential to generate as well as absorb acoustic energy in the low frequency region. The impedance predicted by this model shows the importance of capturing the path of the shed vortex. When the vortex path is captured accurately, the impedance predictions agree well with previous experimental and CFD results, for example predicting the potential for generation of acoustic energy at higher frequencies. For "long" holes, a simplified model which combines Howe's model with plane acoustic waves within the hole is developed. It is shown that the most important effect in this case is the acoustic non-compactness of the hole.

Aerodynamic Effects of High Turbulence Intensity on a Variable-Speed Power-Turbine Blade with Large Incidence and Reynolds Number Variations

NASA Technical Reports Server (NTRS)

Flegel, Ashlie B.; Giel, Paul W.; Welch, Gerard E.

2014-01-01

The effects of high inlet turbulence intensity on the aerodynamic performance of a variable speed power turbine blade are examined over large incidence and Reynolds number ranges. These results are compared to previous measurements made in a low turbulence environment. Both high and low turbulence studies were conducted in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. The purpose of the low inlet turbulence study was to examine the transitional flow effects that are anticipated at cruise Reynolds numbers. The current study extends this to LPT-relevant turbulence levels while perhaps sacrificing transitional flow effects. Assessing the effects of turbulence at these large incidence and Reynolds number variations complements the existing database. Downstream total pressure and exit angle data were acquired for 10 incidence angles ranging from +15.8deg to -51.0deg. For each incidence angle, data were obtained at five flow conditions with the exit Reynolds number ranging from 2.12×10(exp 5) to 2.12×10(exp 6) and at a design exit Mach number of 0.72. In order to achieve the lowest Reynolds number, the exit Mach number was reduced to 0.35 due to facility constraints. The inlet turbulence intensity, Tu, was measured using a single-wire hotwire located 0.415 axial-chord upstream of the blade row. The inlet turbulence levels ranged from 8 to 15 percent for the current study. Tu measurements were also made farther upstream so that turbulence decay rates could be calculated as needed for computational inlet boundary conditions. Downstream flow field measurements were obtained using a pneumatic five-hole pitch/yaw probe located in a survey plane 7 percent axial chord aft of the blade trailing edge and covering three blade passages. Blade and endwall static pressures were acquired for each flow condition as well. The blade loading data show that the suction surface separation that was evident at many of the low Tu conditions has been eliminated. At the extreme positive and negative incidence angles, the data show substantial differences in the exit flow field. These differences are attributable to both the higher inlet Tu directly and to the thinner inlet endwall boundary layer that the turbulence grid imposes.

Aerodynamic Effects of High Turbulence Intensity on a Variable-Speed Power-Turbine Blade With Large Incidence and Reynolds Number Variations

NASA Technical Reports Server (NTRS)

Flegel, Ashlie B.; Giel, Paul W.; Welch, Gerard E.

2014-01-01

The effects of high inlet turbulence intensity on the aerodynamic performance of a variable speed power turbine blade are examined over large incidence and Reynolds number ranges. These results are compared to previous measurements made in a low turbulence environment. Both high and low turbulence studies were conducted in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. The purpose of the low inlet turbulence study was to examine the transitional flow effects that are anticipated at cruise Reynolds numbers. The current study extends this to LPT-relevant turbulence levels while perhaps sacrificing transitional flow effects. Assessing the effects of turbulence at these large incidence and Reynolds number variations complements the existing database. Downstream total pressure and exit angle data were acquired for 10 incidence angles ranging from +15.8deg to -51.0deg. For each incidence angle, data were obtained at five flow conditions with the exit Reynolds number ranging from 2.12×10(exp 5) to 2.12×10(exp 6) and at a design exit Mach number of 0.72. In order to achieve the lowest Reynolds number, the exit Mach number was reduced to 0.35 due to facility constraints. The inlet turbulence intensity, Tu, was measured using a single-wire hotwire located 0.415 axial-chord upstream of the blade row. The inlet turbulence levels ranged from 8 to 15 percent for the current study. Tu measurements were also made farther upstream so that turbulence decay rates could be calculated as needed for computational inlet boundary conditions. Downstream flow field measurements were obtained using a pneumatic five-hole pitch/yaw probe located in a survey plane 7 percent axial chord aft of the blade trailing edge and covering three blade passages. Blade and endwall static pressures were acquired for each flow condition as well. The blade loading data show that the suction surface separation that was evident at many of the low Tu conditions has been eliminated. At the extreme positive and negative incidence angles, the data show substantial differences in the exit flow field. These differences are attributable to both the higher inlet Tu directly and to the thinner inlet endwall boundary layer that the turbulence grid imposes.

Fundamental Aeronautics Program: Supersonics Project - Channeled Center-Body Inlet Experiment Overview

NASA Technical Reports Server (NTRS)

St. John, Clint; Ratnayake, Nalin A.; Frederick, Mike

2012-01-01

The presentation describes supersonic flight testing accomplished on a novel mixed-compression axisymmetric inlet utilizing channels for off-design flow matching rather than a translating centerbody concept.

Fundamental Aeronautics Program: Supersonics Project. Channeled Center-Body Inlet Experiment Overview

NASA Technical Reports Server (NTRS)

SaintJohn, Clint; Ratnayake, Nalin; Frederick, Mike

2012-01-01

The presentation describes supersonic flight testing accomplished on a novel mixed compression axisymmetric inlet utilizing channels for off design flow matching rather than a translating centerbody concept.

Stability analysis of BWR nuclear-coupled thermal-hyraulics using a simple model

DOE Office of Scientific and Technical Information (OSTI.GOV)

Karve, A.A.; Rizwan-uddin; Dorning, J.J.

1995-09-01

A simple mathematical model is developed to describe the dynamics of the nuclear-coupled thermal-hydraulics in a boiling water reactor (BWR) core. The model, which incorporates the essential features of neutron kinetics, and single-phase and two-phase thermal-hydraulics, leads to simple dynamical system comprised of a set of nonlinear ordinary differential equations (ODEs). The stability boundary is determined and plotted in the inlet-subcooling-number (enthalpy)/external-reactivity operating parameter plane. The eigenvalues of the Jacobian matrix of the dynamical system also are calculated at various steady-states (fixed points); the results are consistent with those of the direct stability analysis and indicate that a Hopf bifurcationmore » occurs as the stability boundary in the operating parameter plane is crossed. Numerical simulations of the time-dependent, nonlinear ODEs are carried out for selected points in the operating parameter plane to obtain the actual damped and growing oscillations in the neutron number density, the channel inlet flow velocity, and the other phase variables. These indicate that the Hopf bifurcation is subcritical, hence, density wave oscillations with growing amplitude could result from a finite perturbation of the system even where the steady-state is stable. The power-flow map, frequently used by reactor operators during start-up and shut-down operation of a BWR, is mapped to the inlet-subcooling-number/neutron-density (operating-parameter/phase-variable) plane, and then related to the stability boundaries for different fixed inlet velocities corresponding to selected points on the flow-control line. The stability boundaries for different fixed inlet subcooling numbers corresponding to those selected points, are plotted in the neutron-density/inlet-velocity phase variable plane and then the points on the flow-control line are related to their respective stability boundaries in this plane.« less

Three-step cylindrical seal for high-performance turbomachines

NASA Technical Reports Server (NTRS)

Hendricks, Robert C.

1987-01-01

A three-step cylindrical seal configuration representing the seal for a high performance turbopump (e.g., the space shuttle main engine fuel pump) was tested under static (nonrotating) conditions. The test data included critical mass flux and pressure profiles over a wide range of inlet temperatures and pressures for fluid nitrogen and fluid hydrogen with the seal in concentric and fully eccentric positions. The critical mass flux (leakage rate) was 70% that of an equivalent straight cylindrical seal with a correspondingly higher pressure drop based on the same flow areas of 0.3569 sq cm but 85% that of the straight seal based on the third-step flow area of 0.3044 sq cm. The mass flow rates for the three step cylindrical seal in the fully eccentric and concentric positions were essentially the same, and the trends in flow coefficient followed those of a simple axisymmetric inlet configuration. However, for inlet stagnation temperatures less than the thermodynamic critical temperature the pressure profiles exhibited a flat region throughout the third step of the seal, with the pressure magnitude dependent on the inlet stagnation temperature. Such profiles represent an extreme positive direct stiffness. These conditions engendered a crossover in the pressure profile upstream of the postulated choke that resulted in a local negative stiffness. Flat and crossover profiles resulting from choking within the seal are practically unknown to the seal designer. However, they are of critical importance to turbomachine stability and must be integrated into any dynamic analysis of a seal of this configuration. In addition, choking is highly dependent on geometry, inlet-to-backpressure ratio, and inlet temperature and can occur within the seal even though the backpressure is above the critical pressure.

Evaluation of adsorption effects on measurements of ammonia, acetic acid, and methanol

NASA Astrophysics Data System (ADS)

Yokelson, R. J.; Christian, T. J.; Bertschi, I. T.; Hao, W. M.

2003-10-01

We examined how adsorption and desorption of gases from inlets and a cell could affect the accuracy of closed-cell FTIR measurements of carbon dioxide (CO2), carbon monoxide (CO), methane (CH4), nitric oxide (NO), nitrogen dioxide (NO2), methanol (CH3OH), acetic acid (CH3COOH), and ammonia (NH3). When standards were delivered to the cell through a stainless steel inlet, temporarily reduced transmission was observed for CH3OH and NH3. However, a halocarbon wax coated inlet (normally used on the system) had excellent transmission (comparable to room temperature Teflon) for both CH3OH and NH3, even at temperatures as low as 5°C. Thus the wax is valuable for coating sampling system components that cannot be fashioned from Teflon. The instrument had a delayed response (˜10-40 s) for NH3 only, which was attributed to passivation of the Pyrex multipass cell. To determine sampling artifacts that could arise from the complex sample matrix presented by smoke, the closed-cell FTIR system was intercompared with an open-path FTIR system (which is immune to sampling artifacts) in well-mixed smoke. A similar cell passivation delay for NH3 was the only artifact found in this test. Overall, the results suggest that ˜10 s is sufficient to detect >80% of an NH3/CO ratio sampled by our fast-flow, closed-cell system. Longer sampling times or consecutive samples return better results. In field campaigns the closed-cell system sampling times were normally 10 to >100 s so NH3 was probably underestimated by 5-15%.

Evaluation of centrifugal compressor performance with water injection

NASA Technical Reports Server (NTRS)

Beede, William L; Hamrick, Joseph T; Withee, Joseph R , Jr

1951-01-01

The effects of water injection on a compressor are presented. To determine the effects of varying water-air ratio, the compressor was operated at a constant equivalent impeller speed over a range of water-air ratios and weight flows. Operation over a range of weight flows at one water-air ratio and two inlet air temperatures was carried out to obtain an indication of the effects of varying inlet air temperature. Beyond a water-air ratio of 0.03 there was no increase in maximum air-weight flow, a negligible rise in peak total-pressure ratio, and a decrease in peak adiabatic efficiency. An increase in inlet air temperature resulted in an increase in the magnitude of evaporation. An analysis of data indicated that the magnitude of evaporation within the compressor impeller was small.

Injector Element which Maintains a Constant Mean Spray Angle and Optimum Pressure Drop During Throttling by Varying the Geometry of Tangential Inlets

NASA Technical Reports Server (NTRS)

Trinh, Huu P. (Inventor); Myers, William Neill (Inventor)

2014-01-01

A method for determining the optimum inlet geometry of a liquid rocket engine swirl injector includes obtaining a throttleable level phase value, volume flow rate, chamber pressure, liquid propellant density, inlet injector pressure, desired target spray angle and desired target optimum delta pressure value between an inlet and a chamber for a plurality of engine stages. The tangential inlet area for each throttleable stage is calculated. The correlation between the tangential inlet areas and delta pressure values is used to calculate the spring displacement and variable inlet geometry. An injector designed using the method includes a plurality of geometrically calculated tangential inlets in an injection tube; an injection tube cap with a plurality of inlet slots slidably engages the injection tube. A pressure differential across the injector element causes the cap to slide along the injection tube and variably align the inlet slots with the tangential inlets.

Optimization of Tangential Mass Injection for Minimizing Flow Separation in a Scramjet Inlet

DTIC Science & Technology

1991-12-01

34 Aerospace EnQineering, Vol. 11. No. 8, August 1991, p.23. 26. Heppenheimer , Thomas A . Lecture notes from Hypersonic Technologies seminar. University...AFIT/GAE/ENY,/9 lD-2 ( /~ AD-A243 868 "DTIC OPTIMIZATION OF TANGENTIAL MASS INJECTION FOR MINIMIZING FLOW SEPARATION IN A SC.R-.MJET INLET THESIS...OF TANGENTIAL MASS INJECTION FOR MINIMIZING FLOW SEPARATION IN A SCRAMJET INLEr THESIS Presented to the Faculty of the School of E.ngineering of the

Effects of Inlet Distortion on Aeromechanical Stability of a Forward-Swept High-Speed Fan

NASA Technical Reports Server (NTRS)

Herrick, Gregory P.

2011-01-01

Concerns regarding noise, propulsive efficiency, and fuel burn are inspiring aircraft designs wherein the propulsive turbomachines are partially (or fully) embedded within the airframe; such designs present serious concerns with regard to aerodynamic and aeromechanic performance of the compression system in response to inlet distortion. Separately, a forward-swept high-speed fan was developed to address noise concerns of modern podded turbofans; however this fan encounters aeroelastic instability (flutter) as it approaches stall. A three-dimensional, unsteady, Navier-Stokes computational fluid dynamics code is applied to analyze and corroborate fan performance with clean inlet flow. This code, already validated in its application to assess aerodynamic damping of vibrating blades at various flow conditions, is modified and then applied in a computational study to preliminarily assess the effects of inlet distortion on aeroelastic stability of the fan. Computational engineering application and implementation issues are discussed, followed by an investigation into the aeroelastic behavior of the fan with clean and distorted inlets.

The operational stability of a centrifugal compressor and its dependence on the characteristics of the subcomponents

DOE Office of Scientific and Technical Information (OSTI.GOV)

Hunziker, R.; Gyarmathy, G.

1994-04-01

A centrifugal compressor was tested with three different diffusers with circular-arc vanes. The vane inlet angle was varied from 15 to 30 deg. Detailed static wall pressure measurements show that the pressure field in the diffuser inlet is very sensitive to flow rate. The stability limit regularly occurred at the flow rate giving the maximum pressure rise for the overall stage. Mild surge arises as a dynamic instability of the compression system. The analysis of the pressure rise characteristic of each individual subcomponent (impeller, diffuser inlet, diffuser channel,...) reveals their contribution to the overall pressure rise. The diffuser channels playmore » an inherently destabilizing role while the impeller and the diffuser inlet are typically stabilizing. The stability limit was mainly determined by a change in the characteristic of the diffuser inlet. Further, the stability limit was found to be independent of the development of inducer-tip recirculation.« less

Performance Prediction of Darrieus-Type Hydroturbine with Inlet Nozzle Operated in Open Water Channels

NASA Astrophysics Data System (ADS)

Nakashima, K.; Watanabe, S.; Matsushita, D.; Tsuda, S.; Furukawa, A.

2016-11-01

Small hydropower is one of the renewable energies and is expected to be effectively used for local supply of electricity. We have developed Darrieus-type hydro-turbine systems, and among them, the Darrieus-turbine with a weir and a nozzle installed upstream of turbine is, so far, in success to obtain more output power by gathering all water into the turbine. However, there can several cases exist, in which installing the weir covering all the flow channel width is unrealistic, and in such cases, the turbine should be put alone in open channels without upstream weir. Since the output power is very small in such a utilization of small hydropower, it is important to derive more power for the cost reduction. In the present study, we parametrically investigate the preferable shape of the inlet nozzle for the Darrieus-type hydroturbine operated in an open flow channel. Experimental investigation is carried out in the open channel in our lab. Tested inlet nozzles are composed of two flat plates with the various nozzle converging angles and nozzle outlet (runner inlet) widths with the nozzle inlet width kept constant. As a result, the turbine with the nozzles having large converging angle and wide outlet width generates higher power. Two-dimensional unsteady numerical simulation is also carried out to qualitatively understand the flow mechanism leading to the better performance of turbine. Since the depth, the width and the flow rate in the real open flow channels are different from place to place and, in some cases from time to time, it is also important to predict the onsite performance of the hydroturbine from the lab experiment at planning stage. One-dimensional stream-tube model is developed for this purpose, in which the Darrieus-type hydroturbine with the inlet nozzle is considered as an actuator-disk modelled based on our experimental and numerical results.

The Prominent Role of the Upstream Conditions on the Large-scale Motions of a Turbulent Channel Flow

NASA Astrophysics Data System (ADS)

Castillo, Luciano; Dharmarathne, Suranga; Tutkun, Murat; Hutchins, Nicholas

2017-11-01

In this study we investigate how upstream perturbations in a turbulent channel flow impact the downstream flow evolution, especially the large-scale motions. Direct numerical simulations were carried out at a friction Reynolds number, Reτ = 394 . Spanwise varying inlet blowing perturbations were imposed at 1 πh from the inlet. The flow field is decomposed into its constituent scales using proper orthogonal decomposition. The large-scale motions and the small-scale motions of the flow field are separated at a cut-off mode number, Mc. The cut-off mode number is defined as the number of the mode at which the fraction of energy recovered is 55 % . It is found that Reynolds stresses are increased due to blowing perturbations and large-scale motions are responsible for more than 70 % of the increase of the streamwise component of Reynolds normal stress. Surprisingly, 90 % of Reynolds shear stress is due to the energy augmentation of large-scale motions. It is shown that inlet perturbations impact the downstream flow by means of the LSM.

Digital integrated control of a Mach 2.5 mixed-compression supersonic inlet and an augmented mixed-flow turbofan engine

NASA Technical Reports Server (NTRS)

Batterton, P. G.; Arpasi, D. J.; Baumbick, R. J.

1974-01-01

A digitally implemented integrated inlet-engine control system was designed and tested on a mixed-compression, axisymmetric, Mach 2.5, supersonic inlet with 45 percent internal supersonic area contraction and a TF30-P-3 augmented turbofan engine. The control matched engine airflow to available inlet airflow. By monitoring inlet terminal shock position and over-board bypass door command, the control adjusted engine speed so that in steady state, the shock would be at the desired location and the overboard bypass doors would be closed. During engine-induced transients, such as augmentor light-off and cutoff, the inlet operating point was momentarily changed to a more supercritical point to minimize unstarts. The digital control also provided automatic inlet restart. A variable inlet throat bleed control, based on throat Mach number, provided additional inlet stability margin.

Computational Fluid Dynamics (CFD) Simulation of Hypersonic Turbine-Based Combined-Cycle (TBCC) Inlet Mode Transition

NASA Technical Reports Server (NTRS)

Slater, John W.; Saunders, John D.

2010-01-01

Methods of computational fluid dynamics were applied to simulate the aerodynamics within the turbine flowpath of a turbine-based combined-cycle propulsion system during inlet mode transition at Mach 4. Inlet mode transition involved the rotation of a splitter cowl to close the turbine flowpath to allow the full operation of a parallel dual-mode ramjet/scramjet flowpath. Steady-state simulations were performed at splitter cowl positions of 0deg, -2deg, -4deg, and -5.7deg, at which the turbine flowpath was closed half way. The simulations satisfied one objective of providing a greater understanding of the flow during inlet mode transition. Comparisons of the simulation results with wind-tunnel test data addressed another objective of assessing the applicability of the simulation methods for simulating inlet mode transition. The simulations showed that inlet mode transition could occur in a stable manner and that accurate modeling of the interactions among the shock waves, boundary layers, and porous bleed regions was critical for evaluating the inlet static and total pressures, bleed flow rates, and bleed plenum pressures. The simulations compared well with some of the wind-tunnel data, but uncertainties in both the windtunnel data and simulations prevented a formal evaluation of the accuracy of the simulation methods.

Open-loop heat-recovery dryer

DOEpatents

TeGrotenhuis, Ward Evan

2013-11-05

A drying apparatus is disclosed that includes a drum and an open-loop airflow pathway originating at an ambient air inlet, passing through the drum, and terminating at an exhaust outlet. A passive heat exchanger is included for passively transferring heat from air flowing from the drum toward the exhaust outlet to air flowing from the ambient air inlet toward the drum. A heat pump is also included for actively transferring heat from air flowing from the passive heat exchanger toward the exhaust outlet to air flowing from the passive heat exchanger toward the drum. A heating element is also included for further heating air flowing from the heat pump toward the drum.

Tilt Nacelle Vertical and Short Takeoff and Landing Engine

NASA Image and Video Library

1979-03-21

Center Director John McCarthy, left, and researcher Al Johns pose with a one-third scale model of a Grumman Aerospace tilt engine nacelle for Vertical and Short Takeoff and Landing (V/STOL) in the 9- by 15-Foot Low Speed Wind Tunnel at the National Aeronautics and Space Administration (NASA) Lewis Research Center. Lewis researchers had been studying tilt nacelle and inlet issues for several years. One area of concern was the inlet flow separation during the transition from horizontal to vertical flight. The separation of air flow from the inlet’s internal components could significantly stress the fan blades or cause a loss of thrust. In 1978 NASA researchers Robert Williams and Al Johns teamed with Grumman’s H.C. Potonides to develop a series of tests in the Lewis 9- by 15-foot tunnel to study a device designed to delay the flow separation by blowing additional air into the inlet. A jet of air, supplied through the hose on the right, was blown over the inlet surfaces. The researchers verified that the air jet slowed the flow separation. They found that the blowing on boundary layer control resulted in a doubling of the angle-of-attack and decreases in compressor blade stresses and fan distortion. The tests were the first time the concept of blowing air for boundary layer control was demonstrated. Boundary layer control devices like this could result in smaller and lighter V/STOL inlets.

Design of Single Stage Axial Turbine with Constant Nozzle Angle Blading for Small Turbojet

NASA Astrophysics Data System (ADS)

Putra Adnan, F.; Hartono, Firman

2018-04-01

In this paper, an aerodynamic design of a single stage gas generator axial turbine for small turbojet engine is explained. As per design requirement, the turbine should be able to deliver power output of 155 kW at 0.8139 kg/s gas mass flow, inlet total temperature of 1200 K and inlet total pressure of 335330 Pa. The design phase consist of several steps, i.e.: determination of velocity triangles in 2D plane, 2D blading design and 3D flow analysis at design point using Computational Fluid Dynamics method. In the determination of velocity triangles, two conditions are applied: zero inlet swirl (i.e. the gas flow enter the turbine at axial direction) and constant nozzle angle design (i.e. the inlet and outlet angle of the nozzle blade are constant from root to tip). The 2D approach in cascade plane is used to specify airfoil type at root, mean and tip of the blade based on inlet and outlet flow conditions. The 3D approach is done by simulating the turbine in full configuration to evaluate the overall performance of the turbine. The observed parameters including axial gap, stagger angle, and tip clearance affect its output power. Based on analysis results, axial gap and stagger angle are positively correlated with output power up to a certain point at which the power decreases. Tip clearance, however, gives inversely correlation with output power.

Numerical simulation of tsunami generation by cold volcanic mass flows at Augustine Volcano, Alaska

USGS Publications Warehouse

Waythomas, C.F.; Watts, P.; Walder, J.S.

2006-01-01

Many of the world's active volcanoes are situated on or near coastlines. During eruptions, diverse geophysical mass flows, including pyroclastic flows, debris avalanches, and lahars, can deliver large volumes of unconsolidated debris to the ocean in a short period of time and thereby generate tsunamis. Deposits of both hot and cold volcanic mass flows produced by eruptions of Aleutian arc volcanoes are exposed at many locations along the coastlines of the Bering Sea, North Pacific Ocean, and Cook Inlet, indicating that the flows entered the sea and in some cases may have initiated tsunamis. We evaluate the process of tsunami generation by cold granular subaerial volcanic mass flows using examples from Augustine Volcano in southern Cook Inlet. Augustine Volcano is the most historically active volcano in the Cook Inlet region, and future eruptions, should they lead to debris-avalanche formation and tsunami generation, could be hazardous to some coastal areas. Geological investigations at Augustine Volcano suggest that as many as 12-14 debris avalanches have reached the sea in the last 2000 years, and a debris avalanche emplaced during an A.D. 1883 eruption may have initiated a tsunami that was observed about 80 km east of the volcano at the village of English Bay (Nanwalek) on the coast of the southern Kenai Peninsula. Numerical simulation of mass-flow motion, tsunami generation, propagation, and inundation for Augustine Volcano indicate only modest wave generation by volcanic mass flows and localized wave effects. However, for east-directed mass flows entering Cook Inlet, tsunamis are capable of reaching the more populated coastlines of the southwestern Kenai Peninsula, where maximum water amplitudes of several meters are possible.

Liners for ion transport membrane systems

DOEpatents

Carolan, Michael Francis; Miller, Christopher Francis

2010-08-10

Ion transport membrane system comprising (a) a pressure vessel comprising an interior, an exterior, an inlet, an inlet conduit, an outlet, and an outlet conduit; (b) a plurality of planar ion transport membrane modules disposed in the interior of the pressure vessel and arranged in series, each membrane module comprising mixed metal oxide ceramic material and having an interior region and an exterior region, wherein the inlet and the outlet of the pressure vessel are in flow communication with exterior regions of the membrane modules; (c) a gas manifold having an interior surface wherein the gas manifold is in flow communication with the interior region of each of the planar ion transport membrane modules and with the exterior of the pressure vessel; and (d) a liner disposed within any of the inlet conduit, the outlet conduit, and the interior surface of the gas manifold.